Design proposal FB-24 CD Report Final issue version

CRANFIELD UNIVERSITY

SCHOOL OF ENGINEERING

MSc THESIS

Academic years 2003 – 2006

GEOFFREY A WARDLE MSc CEng

ADVANCED INTERDICTION AIRCRAFT SYSTEM

CONCEPTUAL DESIGN STUDY (PROJECT NOVA).

INDIVIDUAL RESEARCH PROJECT

Supervisor: Professor J. P. FIELDING

February 2006

This thesis is submitted in partial (30%) fulfilment of the requirements for the degree of Master of

Science in Aircraft Engineering.

Cranfield University 2006. All rights reserved. No part of this publication may be reproduced

without the written permission of the copyright holder.

AIA SYSTEM PROJECT NOVA. BAE Systems / Cranfield University.

G. A. Wardle MSc CEng. MSc Individual Research Project Thesis

i

Theses “Health” Warning

This thesis has been assessed as of satisfactory standard for the award of a Master of

Science degree in Aircraft Engineering. This thesis covers part of the assessment

concerned with the Individual Research Project. Readers must be aware that the work

contained is not necessarily 100% correct, and caution should be exercised if this

thesis or the data it contains is being used for future work. If in doubt, please refer to

the supervisor named in the thesis, or the Aerospace Engineering Group.

All of the views and material contained within this document are the sole research of

the author and are not meant to directly imply the intentions of the Joint Strike Fighter

Project Office, National Security Agency or any contractor, or any third party at this

date. Although the USAF awarded contracts for studies into extending the combat

range and enhancing the capabilities of both the F-35 and the F/A-22 in 2004 this is

not representative of the results or based on the results of any part of that body of

research which is secret.

This thesis is an unsolicited conceptual design study, which has been reviewed by Mr

Robert A. Ruszkowski, Jr Senior Staff Engineer of Lockheed Martin ADP for whose

advice I am eternally grateful.

This document contains no material governed by ITAR restrictions and the

distribution of all information contained within this document is unlimited public

release and has been approval by: - US D o D, UK M o D representatives, Lockheed

Martin (JSF Programme) and BAE Systems (Future Offensive Air Systems). This

document and any part thereof cannot be reproduced by any means for distribution

without written permission form Cranfield University and the author.

None of the data contained within this project is to be used in any format or for any

other research project without consultation with Cranfield University School of

Engineering supervisor named on the front cover.

Acknowledgements

Mr Robert A. Ruszkowski, Jr Senior Staff Engineer at Lockheed Martin ADP for

whose advice I am eternally grateful, Andy Bruce Design Lead for F-35C empennage

for private help with background material on JAST and proof reading the final draft,

and Phil Read JSF BAESYSTEMS security officer for obtaining security clearance

for this project, and finally acknowledgement to the following: - Steven A. Brandt:

John J. Bertin: Randall J. Stiles: and Ray. Whitford: of the UASFA for producing the

AeroDYNAMIC V2.08 analysis software and the AeroDYNAMIC V3 design

resource CD which was a major aid in producing this thesis.



ii

Abstract.

The objective of this thesis was to produce a conceptual design study for a force

package consisting of a two seat advanced interdiction aircraft and a complementary

UCAV capable of replacing the UK Royal Air Force Tornado, Royal Australian Air

Force F-111, and USAF F-117 and F-15E air assets, both of these aircraft would have

greatly enhanced capabilities in stealth, range, and supercruise capability. The

performance being measured against a representative future mission profile produced

by the American Institute of Aeronautics and Astronautics, and the USAF Academy.

These aircraft represent a fusion of F/A-22A, F-35C, and YF-23 technologies with the

innovative F-120 Variable Cycle Engine (flight proven in the Advanced Tactical

Fighter CDA program), to produce single engine interdictor, with a predicted

supercruise capability, and 900mile combat radius, which was capable of carrying

typical current and near future internal weapons loads, for an estimated unit cost of

$75 million in 2006 dollars. The capability to supercruise combined with a high

degree of stealth was reasoned to severely degrade the capability of enemy defences

reducing their response time to a grater extent than that achievable with F-117 stealth

fighter enabling the aircraft to prosecute a high altitude attack in a greatly reduced

threat envelope, clear of AAA, and MANPADS, which have the fastest response

times, compared with the much larger SV300 and SV400 class weapons.

The initial starting point for these aircraft was the current F-35C USN aircraft carrier

variant of the Joint Strike Fighter, with all naval capability removed, and originally a

growth aircraft with a new wing was considered, using a similar development

methodology as the growth of the F/A-18C Hornet into F/A-18E Super Hornet or F-

16C into the YF-16XL. The requirement for supercruise capability over a large

portion of the mission lead to very major airframe changes: - to reduce wave drag by

increasing the finesse ratio of the fuselage, changing the wing planform, increasing

the wing leading edge sweep angle to keep the wing inside the shock cone produced

by the fuselage, and adoption of duel function twin ruddervators eliminating the

horizontal tail surfaces. The resulting aircraft were an evolution of the F-35C into a

larger aircraft suitable for proposed mission, and although much airframe

commonality was lost provision was made to integrate common systems within the

common airframe of both manned and unmanned variants and these systems represent

a greater percentage of overall weapon system cost than airframe itself.

This conceptual design study was for the modification of an exiting airframe design

and was different to that of a clean sheet of paper design consisting of two phases, the

first of which was configuration design and parametric analysis using both classical

analysis and the Jet306 / AeroDYNAMIC V2.08 analysis tool set, and the second was

major structural component layout of the airframe initial structure with systems

integration (the original intention of using hand calculations with PATRAN /

NASTRAN FEA modelling for structural sizing proved impractical due to time and

training resource constraints which also precluded CFD analysis in phase one). This

final design study for both versions of the AIA aircraft contained herein consists of

parametric analysis, initial optimisation and structural layout and constitutes a

feasibility study proposal to meet the requirements. Recommended further work on

the proposed designs includes CFD, and FEA analysis of the drag, aero and structural

loads.



iii

Contents Page

Theses “Health” Warning i

Acknowledgements i

Abstract ii

Contents iii

Figures v

Table‟s xiii

Performance and design analysis charts xiv

Glossary xv

1.0 Introduction 1

1.1. Project brief and SOW 3

2.0 Requirements capture 6

2.1 Current F-35 family of aircraft data 10

2.2 Threat analysis 11

2.2.1 Air threats 11

2.2.2 Surface to air threats 13

2.2.3 Stealth requirements 15

3.0 FB – 24 configuration selection and optimisation 16

3.1 Initial conceptual design studies 16

3.1.1 FB - 24 and F-35C common features 20

3.2 FB – 24 and A-24 Configuration concepts 21

3.2.1 Configuration design challenges 36

3.2.2 Fuselage configuration design selection 37

3.2.3 Wing and Empennage configuration design selection 45

3.2.4 Common FB – 24 and A-24 configuration initial sizing 81



iv

3.2.5 Common FB – 24 and A-24 supersonic drag analysis 128

3.3 Configuration optimisation by parametric analysis 133

3.3.1 Analysis methodology 133

3.3.2 Analysis of NB1 and NB2 on AIA sizing mission 134

3.3 3 Comparison of the results for NB-1 and NB-2 144

4.0 Structural layout and major system integration 150

4.1 Undercarriage integration 151

4.2 Aircrew integration / AI integration 155

4.3 Propulsion system integration 160

4.4 Weapons systems integration 169

4.5 Structural Layout 176

4.6 Fuel tank integration 198

4.7 Aircraft OML G.A. drawings 199

5.0 Conclusions and recommendations for further work 202

References 205

Appendices A: - Detailed primary threat data 208

Appendices B: - Signature reduction 213

Appendices C: - DSI Overview 225

Appendices D: - F-35 Family Overview 227

Appendices E Design supplement 232

E – 1.1:- Structural amendments and concept completion 232

E – 1.2:- Structural weight estimation as modelled 244

E – 2:- Supersonic range and endurance analysis NB2 248

E – 3:- AeroDYNAMIC V 2.08 analysis methodology 253

E – 4:- Comparison of F-24 and A-24 with the RFP targets 265



v

Figures Page

1:- Conceptual strike aircraft RCS pole model 2

2:- UASF Strike Airframe life estimates 3

3:- Statement of work flow chart 4

4:- Jet306 / AeroDYNAMIC analysis examples 5

5:- AIA FB-24 mission profile 8

6:- The MAPO MiG-29M: tactical fighter 12

7:- The MAPO MiG-31B interceptor fighter 12

8:- The Il A-50 AWAC aircraft 13

9:- The S-300V (SA-12) missile system 14

10:- The QFD House of Quality process 18

11:- First tier House of Quality of the AIA 19

12:- Internal weapon options for FB-24 21

13:- Typical performance map for a tactical fighter aircraft 23

14:- General flow regimes encountered by tactical fighters 23

15:- Demonstration wing for twist and camber research NASA 25

16:- Effects of twist and camber on longitudinal aerodynamics 26

17:- Northrop / MDA YF-23 showing ruddervator configuration 27

18:- Boeing X-45 UCAV demonstrator showing tailless configuration 28

19:- Lockheed Martin X-44 showing 2-D vectored tailless configuration 29

20:- F-35A forward fuselage showing extent of the cockpit 32

21:- F/A-22A ACES II ejection seat showing installation angle 32

22:- F-22B Twin seat trainer variant of F/A-22 not built 33

23:- Thrust to weight-v-Wing loading for strike aircraft 35

24:- YF-22 Cross sectional area plot using Jet 306 showing intake effects 37



vi

25:- Wing position on combat aircraft fuselages 39

26:- Advantages of the shoulder mounted wing location 41

27:- Effects of area – rule shaping on supersonic aircraft 43

28:- Flow fields around cylindrical and chined bodies 44

29:- Evolution of wing planforms for combat aircraft 46

30:- YF-22 and Eurofighter configuration comparison 47

31:- Pressure and shear forces on an airfoil 43

32:- Nomenclature for aerodynamic forces in the pitch plane 43

33:- Moment balance to trim an aircraft 44

34:- Aircraft reference axes and corresponding aerodynamic moments 45

35:- Characteristics of airfoil sections 53

36:- Airfoil forces and moments 54

37:- Airfoil centre of pressure 54

38:- Aerodynamic centre 1 influence of location on moment 55

39:- Aerodynamic centre 2 location where moment is independent of alpha 55

40:- Special airfoil profiles 57

41:- Types of trailing edge flaps 60

42:- Lift and drag coefficient curves for wings with trailing edge flaps 61

43:- Types of boundary layed control devices 61

44:- Effects of boundary layer control devices on the lift curve 62

45:- NACA 0006 AIA wing root airfoil analysis 63

46:- NACA 64-006 AIA wing tip airfoil analysis 64

47:- Reynolds number effects on lift and drag curves 65

48:- Wing planform RCS spike effects 67

49:- Wing planform influence on design 68



vii

50:- Fuselage edge alignment influence on design 68

51:- Surface current scattering 69

52:- Reasoning behind wing configuration (B) 71

53:- Wing geometry 71

54:- Effect of aspect ratio on lift 72

55:- Wing relationship to aircraft Centre of Gravity 74

56:- The effect of sweep angle on MCRIT for a non tapered wing 76

57:- Initial definitions tail sizing 79

58:- F-35C port side tail group showing VT/rudder and flipper HT 80

59:- YF-23 Ruddervator starboard side, showing relative size and shape 80

60:- The effect of reducing wing loading on instantaneous turn speed 87

61:- Initial wing option (A) YF-22 based planform 88

62:- Initial wing option (B) JAST based planform 88

63:- Initial wing option (C) F-117A based planform 89

64:- Determination of the MAC for the initial option (A) wing 90

65:- Determination of the MAC for the option (B) wing 91

66:- Determination of the MAC for the option (C) wing 93

67:- Determination of the baseline F-35C MAC wing 94

68:- Original option (B) wing on FB-24 fuselage with F-35C sized tails 95

69:- Determination of the MAC for the revised option (A) wing 96

70:- Determination of the MAC for the revised option (B) wing 97

71:- Determination of the MAC for the revised option (C) wing 99

72:- Wing positioning on the fuselage option (A) wing 101

73:- Wing positioning on the fuselage option (B) wing 102

74:- Wing positioning on the fuselage option (C) wing 103



viii

75:- Empennage location and sizing conventional tail (A) wing (NB1) 111

76:- Empennage location and sizing ruddervator tail (A) wing (NB2) 111

77:- Empennage location and sizing conventional tail (B) wing (NB3) 112

78:- Empennage location and sizing ruddervator tail (B) wing (NB4) 112

79:- Empennage location and sizing conventional tail (C) wing (NB5) 113

80:- Empennage location and sizing ruddervator tail (C) wing (NB6) 113

81:- Baseline F-35C conventional tail sizing 114

82:- Large EHA horizontal tail and ruddervator actuator 117

83:- Small EHA rudder and wing trailing edge control surface actuator 117

84:- Multi segment leading edge lap actuator 120

85:- Flight control surface trade study for NB1 / NB2 wing 120

86:- FB-24 (NB1) Flight control surface layout 121






92:- Fuselage Design Breakdown for Jet analysis 134

93:- Surface model cuts on NB1 OML in isometric view 135

94:- Representative F-136-2 engine modelled in CATIA V5 135

95:- Area distribution for NB1 136

96:- Area distribution for NB2 140

97: Design limitations of the current F-35 aircraft 149

98:- Comparison of F-35 CTOL with F/A-22A 149

99:- F/A-22A internal structural layout illustrating complexity 150



ix

100:- FB-24 NB1 Tip back angle analysis 151

101:- FB-24 NB1 Overturn angle analysis 151

102:- FB-24 NB2 Tip back angle analysis 152

103:- FB-24 NB2 Overturn angle analysis 152

104:- Undercarriage retraction common for FB-24 / A-24 154

105:- Undercarriage storage common for FB-24 / A-24 155

106:- Design eye view for aircraft pilot commander FB-24 156

107:- Original crew station layout based on F-22B 157

108:- F-35 Canopy opening arrangement 158

109:- FB-24 Twin crew station integration 158

110:- The A-24 AI unit integration in fwd fuselage 159

111:- PW F-135 cutaway JSF CDA engine 160

112:- PW F-136 undergoing final inspection 161

113:- GE F-120 Schematic illustrating VCE operation modes 161

114:- Longitudinal section showing main components of the F-120 162

115:- Engine integration common to both FB-24 and A-24 164

116:- Common FB-24 / A-24 Intake duct 165

117:- Engine installation 166

118:- Engine airframe mounting 166

119:- Engine transporter 167

120:- Engine mounting and load transfer 168

121:- JASSM on transport mount 170

122:- Installation of SDB‟s 171

123:- Proposed ASRAAM installation 172

124:- Standard F-35 weapons bay with 2,000lb JDAM 173



x

125:- FB-24 / A-24 Weapons bay fit study 173

126:- Final FB-24 / A-24 weapons bay integration 174

127:- Future weapons for AIA the LOAAS 175

128:- Future weapons for AIA the SMACM 175

129:- F-35 Continuous wing substructure 177

130:- F-35 Fibre placed continuous wing skin 177

131:- F/A-22A Two separate wing substructure 178

132:- F/A-22A Wing skin / substructure assembly 178

133:-FB-24 / A-24 Phase 1 wing layout 179

134:- Phase 1 wing detailed description 180

135:- FB-24 / A-24 Phase 2 wing layout 181

136:- Phase 2 wing detailed description 182

137:- FB-24 / A-24 Final wing layout 183

138:- Final wing layout detailed description 184

139:- Detailed structural layout of FB-24 forward fuselage top view 186

140:- Detailed structural layout of FB-24 forward fuselage underside view 187

141:- Forward fuselage additional principle structure dimensions 188

142:- Detailed structural layout of centre fuselage top view 189

143:- Detailed structure layout of centre fuselage underside view 190

144:- Centre fuselage and wing integration 191

145:- Differences in centre fuselage frame between FB-24 and F-35 191

146:- Detailed aft fuselage layout top view 192

147:- Detailed aft fuselage underside view 193

148:- Ruddervator structural layout 194

149:- A-24 AI integration into complete forward fuselage structure 194



xi

150:- Complete airframe structural integration model 195

151:- F-35C plan view for comparison with figure 149 196

152:- F-35C side view for comparison with figure 149 196

153:- FB-24 Fuel tank integration model 197

154:- A-24 Fuel tank integration model 197

155:- Dimensioned FB-24 GA drawing Plan view 199

156:- Dimensioned FB-24 GA drawing Side view 200

157:- Dimensioned FB-24 GA drawing Front view 201

A.1:- MiG Light Weight Fighter Project aircraft concept 208

A.2:- Chinese J-10 Light Weight point defence fighter 208

B.1:- Tornado GR4 current RAF deep strike aircraft 215

B.2:- USAF F-15E Strike Eagle 216

B.3:- USAF F-111E similar to the RAAF F-111C‟s 216

B.4:- Spike alignment or plan from alignment on the F-35C 217

B.5:- Spike side lobe alignment on the F-35C 218

B.6:- F-117A illustrating facetted approach to RCS reduction 219

B.7:- The F-35 SigMA pole model about to undergo RCS measurement 220

B.8:- The F-35 SigMA model mounted at the test range 221

B.9:- Pratt & Whitney LO nozzle under ground based test 223

B.10:- General Electric AVEN nozzle (LOAN) on show 223

B.11:- F/A-22A 2-D vectoring nozzle installed on an F/A-22A 224

C.1:- Diverterless Supersonic Intake on CDA mock – up 226

C.2:- GFD Model of the flow fields around the DSI 227

D.1:- The F-35A USAF CTOL aircraft three view 228

D.2:- The X-35A in flight 229



xii

D 3:- The F-35B USMC / RN /RAF aircraft three view 230

D.4:- The X-35B in hover 230

D.5:- The F-35C USN aircraft three view 231

D.6:- The F-35C maximum stores carriage capability 231

E.1:- The FB-24 MOSC hatch structure 233

E.2:- The FB-24 MOSC hatch integration with the forward fuselage 233

E.3:- Revised A-24 forward fuselage layout top view 234

E.4:- Typhoon refuelling indicating required A-24 field of regard 235

E.5:- Revised structural arrangement common aft fuselage top view 236

E.6:- Revised structural arrangement common aft fuselage underside view 237

E.7:- Revised common centre fuselage structure top view 238

E.8:- Revised FB-24 detailed structural layout model 239

E.9:- Revised A-24 detailed structural layout model 240

E.10:- FB-24 / A-24 revised common wing design 241

E.11:- FB-24 / A-24 metallic wing joint philosophy 242

E.12:- FB-24 / A-24 composite wing joint philosophy 242

E.13:- FB-24 / A-24 composite substructure common wing design 241

E.14:- A-24 Evolution general arrangement model 243

E.15:- Structural weight measurement methodology for frames 244

E.16:- Structural weight measurement methodology for keels and longerons 245

E.17:- Structural weight measurement methodology for wing skins 248

E.18:- Structural weight measurement methodology for fuselage skins 248

E.19:- Airplane geometry for downwash prediction 259

E.20:- AeroDYNAMIC V 2.08 Lifting surface analysis 265

E.21:- AeroDYNAMIC V 2.08 Surface area analysis 267



xiii

Tables Page

1:- Mission breakdown for the Advanced Interdiction Aircraft 3

2:- Government Furnished Equipment 34

3:- Predicted variation of MNCRIT with sweep angle 77

4:- Experimental variation of MNCRIT with sweep angle 77

5:- Mass: wing loading: and dry thrust / weight ratios 83

6:- Wing configuration study summary 96

7:- Revised wing sizing study to common wing loading 101

8:- Wing configuration layout 102

9:- Tail sizing results for HT, VT, and RV configurations 109

10:- Exposed tail surface areas 118

11:- Flight control surface sizings relative to the wing area 127

12:- NB1 and NB2 major component weight from AeroDYNAMIC 144

13:- F-24 / A-24 comparison with RFP targets 202

B-1:- Typical Radar Threats 215

E-1:- Wing and Empennage structural component weight table 245

E-2:- Forward fuselage structural component weight table 246

E-3:- Centre and Aft fuselage structural component weight table 247

E-4:- Major component skin weight estimate 249

E-5:- As drawn weight and Jet 2.08 weight prediction comparison 249

E-6:- F-24 / A-24 Definitive weight statement 250

E-7:- F-24 / A-24 Lift curve analysis 266

E-8:- F-24 / A-24 Aerodynamic analysis 266

E-9:- F-24 / A-24 Drag Polar analysis 268

E-10:- F-24 / A-24 CD vs CL analysis 268



xiv

E-11:- F-24 / A-24 Lift over Drag vs CL analysis 269

E-12:- F-24 / A-24 CL ^1.5 over CL analysis 269

Performance and design analysis charts Page

1(a):- NB1 Total predicted drag variation with Mach number 136

1(b):- NB1 Drag Polar 136

2:- NB1 Lift over drag v lift coefficient at (a) Mach 0.85 and (b) Mach 1.5 137

3:- NB1 Lift curve CL v 137

4:- NB1 CD v CL at (a) Mach 0.85 and (b) Mach 1.5 138

5:- NB1 Performance analysis Thrust and Drag v Mach number 138

6(a):- NB-1 V-n Diagram 139

6(b):- NB1 Manoeuvre diagram 139

7:- NB1 Specific Excess Power curves 139

8(a):- NB2 Total predicted drag variation with Mach number 140

8(b):- NB2 Drag Polar 140

9:- NB2 Lift over drag v lift coefficient at (a) Mach 0.85 and (b) Mach 1.5 141

10:- NB2 Lift curve CL v 141

11:- NB2 CD v CL at (a) Mach 0.85 and (b) Mach 1.5 142

12:- NB2 Performance analysis Thrust and Drag v Mach number 142

13(a):- NB2 V-n Diagram 143

13(b):- NB2 Manoeuvre diagram 143

14:- NB2 Specific Excess Power curves 143



xv

Glossary

ALOSNW Air launched Low Observable Stand off Nuclear Weapon

APU Auxiliary Power Unit

AIA Advanced Interdiction Aircraft

AMRAAM Advanced Medium Range Air to Air Missile

ASRAA Advanced Short Range Air to Air Missile

ATF Advanced Tactical Fighter Programme (now F/A-22A)

B Bomber

BAe British Aerospace now BAESYSTEMS

BCA/BCM Best - cruise - altitude / Best - cruise - Mach

BVR Beyond Visual Range

CDA Concept Demonstrator Aircraft

CFD Computational Fluid Dynamics

CTOL Conventional Take-Off and Landing

CV Aircraft suitable for large aircraft carrier operations

DoD Department of Defence

ECM Electronic Counter Measures

EMPW Electro Magnetic Pulse Weapon

F Fighter

F/A Fighter Attack

FB Fighter Bomber

FEA Finite Element Analysis

FOAS Future Offensive Air System now SUAV (E)

FSAV Future Strategic Air Vehicle

FBW Fly By Wire



xvi

FBL Fly By Light

GR Ground attack and Reconnaissance

GE General Electric (aero engines)

IR Infrared (referring to signature)

JAST Joint Advanced Strike Technology now JSF F-35

JSF Joint Strike Fighter

JSOW Joint Stand Off Weapon

LOAN Low Observable Axisymmetric Nozzle

LMTAS Lockheed Martin Tactical Aircraft Systems

MDA McDonnell Douglas Aircraft now Boeing

MoD Ministry of Defence (United Kingdom)

NG Northrop Grumman

PGW Precision Guided Weapons

P&W Pratt & Whitney (aero engines)

PWSC Preferred Weapons System Concept

RAF Royal Air Force

RAAF Royal Australian Air Force

TSFC Thrust Specific Fuel Consumption

SDD System Development and Demonstration

T/W Thrust to Weight ratio

UCAV Unmanned Combat Air Vehicle

USAFA United States Air Force Academy

Wo, We, Wf Aircraft designed takeoff, empty, and fuel weights

W/S Wing loading



1

1. Introduction.

This Individual Research Project (IRP) thesis forms the submission by the author for

the IRP component of the Cranfield University MSc in Aircraft Engineering, and in

no way represents the views of the F-35 Joint Project Office, the US D o D, UK M o

D, or any F-35 / F/A-22 Contractor Company.

The mission specification for this project is based on the American Institute of

Aeronautics and Astronautics annual aircraft design project competition of 2001/02.

The Request for Proposal (RFP) published by the AIAA is based on recent industrial

project work. Hence this competition provided a useful source of realistic mission

requirements and operational data that form the basis of this submission, additional

information from public domain material for a hypothetical Future Strategic Air

Vehicle program FSAV was also used to tailor this aircraft to the Tornado, F-15E, F-

117A, and F-111C replacement roles. (Reference 1: - Aircraft Design Projects for

Engineering Students: by Jenkinson L. R. and Marchman J. F. Published by AIAA

Education Series in 2003. ISBN 1-56347-619-3. Pages: - 208-209, and USAFA course

FSAV study included in Aerodynamic Version 3.0 Software: Published by the USAFA

and released by AIAA Education Series in 2004: ISBN 1-56347-689-4.)

The Tornado GR4 is the UK‟s only deep strike aircraft, and like the previous

generations of strike aircraft the Tornado was tailored to high speed ultra low level

interdiction missions. Both Gulf War‟s and the Balkans conflict have demonstrated

that interdiction missions are extremely high risk at low level, and the need for high

level strikes with stealth aircraft in these recent conflicts has been apparent.

In fact the exportation of former Soviet Union and Russian Federation advanced

technology SAM, AWACS, and AAA weapons makes survivability with PGW, even

on the medium altitude missions of 20,000 ft for all non stealthy platforms

questionable, without extensive air defence degradation.

The only stealth tactical strike aircraft in service with any nation is the F-117A which

by 2020 will be entering its airframe retirement age, the same is true for the RAF

Tornado Interdictor Strike (IDS) aircraft and USAF F-15E Strike Eagle and both latter

aircraft require an extensive reduction in the enemy air defence network to survive

long enough to deliver mission ordinance against aggressor nations which have

advanced air defence capability. This in turn requires a larger expeditionary force

with specialised ECM screening and SEAD attack aircraft, which increases the overall

cost of any major air interdiction operation, and eliminates the element of surprise. In

addition to the conventional interdiction role this aircraft will form the common

airframe for both manned and unmanned components of a mixed fleet, and by the

2020 time frame the UK will be seeking a new strategic weapons platform, when the

Trident SLBMS submarines reach hull retirement age therefore every possibility

exists that a new aircraft will have a strategic deterrents role, with tactical nuclear

weapons, and become the Future Strike Air Vehicle (FSAV) platform of choice.

To study the UK‟s future platform requirements a manned stealth technology

demonstration model was produced as the Replica LO pole model figure 1, and a UK

variant of the Lockheed F-117 was also considered, both of these projects have been

replaced by SUAV (E) a strategic UAV research program.



2

Therefore the need exists for a new aircraft which can effectively deliver conventional

tactical PGW‟s or compact stand – off nuclear weapons, over long ranges and which

can rapidly deploy with minimum support to regional conflicts world-wide, and

survive in a heavily defended enemy air space without additional air assets.

Improved threat capabilities posed by the S-300V Gladiator and S-400V Triumf SAM

systems dictate that this new aircraft should have the following attributes: - an RCS of

-20db‟s over a frequency range of 0 to 20 GHz : an IR signature at military power

below a wavelength of 8 microns: the capability to fly at high altitude up to 45,000ft

for sustained periods at high Mach number M1.6 to reduce the defence response time:

and the capability to fly supersonically with reduced emissions: and have long range

on internal fuel, which will allow the aircraft to respond to crises around the world

from out of theatre bases which cannot be used by current strike assets, without

carrying external fuel tanks or air to air refuelling.

Approximately 300 aircraft could be required by the USAF for F-15E / F-117A

replacement, with 200 required for the UK RAF to meet a projected FSAV

requirement, and an additional 50 aircraft for the RAAF as F-111C replacements.

Figure 1: - In support of UK future platform solutions full – scale RCS

qualification model was produced which had attributes of a stealthy aircraft in

terms of outer mould lines and plan - form alignment. Source: - BAE Systems

Warton.



3

AERO 481

Figure 2: - USAF Strike airframes that require replacement by 2020. Source: -

AIAA /USAFA Aerodynamic V3.0 included USAFA AERO 481 lecture material.

1.1 Project brief and SOW.

The objective of this project is to produce a conceptual design and structural layout

for both aircraft of the mixed fleet concept, capable of meeting the requirements of

interdiction and strike missions using both conventional and nuclear weapons outlined

above and detailed in section 2. This aircraft will be based on a fusion of F-35C and

F/A-22A technologies.

This project will not use any material which is not within the public domain. The

designations proposed for this aircraft namely FB-24 for the USAF and AIA-1 for the

RAF and RAAF distinguishes this project from any exiting or near future BAE

SYSTEMS / Lockheed Martin Tactical Aircraft projects and prevents confusion with

other programmes of research.

To meet the AIAA and hypothetical FSAV amended mission requirements the F-35C

is to be used as the starting point for this study, with the fuselage extended to give a

fineness ratio comparable to the F/A-22 and a new wing based on one of the

following configurations: - (1) a new wing based on the F/A-22A planform of

increased size as a conservative option: (2) a derivative version of the JAST cranked

arrow head layout , and as the most radical solution (3) a double delta wing similar in

plan form to the F-16XL, all options would have the stretched fuselage and a reduced

fuselage cross section area. The scope and deliverables for this study are shown in the

flow chart figure 3.



4

The maximum GTOW of 71,018lbs (the maximum overload F-35C weight) must not

exceed for the FB-24 in the clean condition (no external stores provision). The

maximum empty weight must not exceed 38,380lbs 18.76% greater than the F-35C,

conceptual weight target breakdown will be as follows:- Pilots plus one M61A

120mm cannon = 2%: engine / nozzle / oils / fluids = 7% (based on the dry weight of

the F110-GE-132 engine which is 5.7% of the GTOW): expendable weapons = 8%:

Structural / avionics / undercarriage / systems = 45% and fuel = 38%. The with a

43,000lb thrust engine the thrust to weight ratio would be 1:0.6 comparable with the

Tornado GR-4, F-15E and the F-117A.

CranfieldUNIVERSITY BAE SYSTEMSDocument No: - WSL2004-IRP PRES 9.

STAGE 1

Configuration generation.

Optimisation analysis and trade study.

Basepoint configuration

selection.

Interim report Sept

2005.

OML Freeze.

Interim oral

report Jan

2006.

STAGE 2

Structural layout and systems integration.

Structural layouts and

trade studies.

Layout Freeze. Systems

integration.

FINAL REPORT.

Feb2006

Figure 16 :- Advanced Interdiction Aircraft Design study flow chart .

Figure 3: - SOW Flow Chart of the FB-24 conceptual design study.

STAGE 1: -

The F-35C public domain data was used to produce the surface model of the F-35C –

230-5 baseline OML configuration to obtain aerodynamic, performance, and

endurance data against the AIAA / FSAV mission requirement. Five proposed aircraft

configurations were then evaluated against the F-35C figures for the same AIAA /

FSAV mission profile to compare, each configuration. The Jet306 Parametric analysis

toolset which is the core of the USAF Academy Aerodynamic V3.0 software was

used for these configuration trade studies.

This tool performs whole aircraft analysis for: - lift: parasite drag: induced drag

supersonic drag: propulsion analysis: weight prediction: constraint analysis: sizing:

optimisation: performance analysis and cost analysis, based on geometry for the

conceptual design and mission profile entered on spreadsheets by the designer with

the fidelity of the analysis depending on the number of section cuts and geometry data

points used this forms the stage 1 trade study. Examples of Jet306 analysis of a test F-

16 are shown in figure 4.



5

0 10 20 30 40 50

Constraints

0

1

2

3

4

5

6

0 20

40

60

80

100

120

140

160

Wing Loading, Wto/S, psf

Th

rust

Lo

ad

ing

, T

sl/W

to

M xM ach

M idTurn

Ps

Ingress

LoTurn

HiTurn

SloTurn

HiCruise

Takeoff

Landing

Actual

Design Point

Desired

- 4 0

- 3 0

- 2 0

- 10

0

10

2 0

3 0

4 0

5 0

0 10 2 0 3 0 4 0 5 0 6 0

Figure 4: Jet306 / AeroDYNAMIC analysis output examples for an F-16 model.

Measures of Merit

The results of the Jet 306 of the five aircraft configurations were then evaluated using

the following measures of merit to down select a single configuration to meet the AIA

/ FSAV mission requirements as detailed in section 2: -

1.1 Weight summary (GTOW, We, Wf, W/S, T/W, Wf /W) using only

internal fuel.

1.2 Aircraft geometry (wing and control surface area, fuselage size and

volume, frontal cross sectional area distribution, wetted area, inlet

and diffuser, landing gear, weapons carriage, sensor and avionics

locations, crew stations, etc.)

1.3 Mission duration, radius or range, fuel burn by mission segment for

each design mission.

1.4 Take-off and landing distance for each design mission including

standard day and icy runway balanced field length at sea level.

1.5 Performance at maneuver weight with 50% internal fuel and with

two ASRAAM missiles and two 2000lb JDAM design mission

loadings.



6

1.5.1 Maximum Mach number at 45,000 ft.

1.5.2 1-g Maximum Thrust Specific Excess Power Envelope

1.5.3 2-g Maximum Thrust Specific Excess Power Envelope

1.5.4 Maximum Thrust Sustained Load Factor Envelope

1.5.5 Maximum Thrust Maneuvering Performance Diagrams

1.5.5.1 34,000 ft

1.5.5.2 45,000 ft

1.6 Flyaway and total life costs.

STAGE 2: -

The down selected configuration was then developed into full CATIA V5 R10 model

for structural analysis and systems integration and this forms the stage 2 design

analysis.

Basepoint structural layouts of the aircraft to meet the systems integration

requirements, for weapons, aircrew, AI system, fuel, and undercarriage were

produced in CATIA V5 R10, and this model was then to be analysed using PATRAN

/ NASTRAN Finite Element Analysis package, however time constraints precluded

this and the structural analysis will differed for a later work package to be perused at a

latter date. This structural analysis will enable initial sizing of the wing substructure

and concluded the conceptual structural layout design freeze.

The final report consists of a complete conceptual design analysis and structural

component layout of the FB-24 Advanced Interdiction Aircraft, although without the

supporting structural analysis, however it forms the basis for the conceptual proposal

submission for the Advanced Interdiction Aircraft as a common manned and

unmanned airframe.

2. Requirements capture.

1. The AIAA and FSAV requirements for the AIA specifies a two-place

advanced deep-interdiction aircraft and ingress and egress portions of the

mission will be flown at supersonic speed at high altitude with target

acquisition and weapons release at high altitude. The mission profile is shown

in figure 5 and broken down in table 1 below, and is flown over a combat

radius of 900 nautical miles. The maximum TOGW of 71,018lbs (the

maximum overload F-35C weight) must not exceed for the FB-24 in the clean

condition (no external stores provision). The maximum empty weight must not

exceed 31,177lbs 3.3% greater than the F-35C.

2. The aircraft must be capable of „all-weather‟ operation including operation

from and on to icy 800 ft runways.

3. The aircraft must operate from forward NATO and allied bases, with the

minimum of support facilities. On these bases the aircraft will be required to

fit into standard NATO shelters.



7

4. The design layout should allow for easy maintenance. Minimum reliance on

support equipment is essential for off-base operations.

5. Structural design limit load factors of +7g to -3g (aircraft clean and with 50

percent internal fuel) are required. An ultimate design factor of 1.4 is to be

applied to reduce weight. The structure must be capable of withstanding a

dynamic pressure (q) of 2133 lb/ft2 (i.e. equivalent to (q) at 800kt) and be

durable and damage tolerant.

6. All fuel tanks must be self-sealing. Aviation fuel to JP-8 and JP-5

specifications is to be assumed.

7. Stability and handling characteristics must meet MIL-F-8785B subsonic

longitudinal static margins to be no grater than +10% and no less than -30%.

8. In addition to the basic mission criteria the design specification requires the

following manoeuvring targets be met (specific excess power, SEP, is defined

as Ps in the following equation: - Ps = [(T/W)-(D/W)] V (where weight = W

=Mg)).

SEP (1g) military thrust (dry), 1.6M at 45,000ft = 0ft/s.

SEP (1g) maximum thrust (wet), 1.6M at 45,000ft = 200ft/s.

SEP (2g) maximum thrust (wet), 1.6M at 45,000ft = 0ft/s.

Maximum instantaneous turn rate, 0.9M at 15,000ft = 8.0degrees/s.

9. The design specification calls for five separate internal weapon capabilities as

shown in figure 4, but the deletion of the external weapons carriage options of

the existing F-35 family:

Two Mk-84 Low Drag GP + two ASRAAM.

Two 2000lb JDAM PIP + two ASRAAM.

Two AGM - 154 JSOW + two ASRAAM.

Six 250lb Boeing Small Diameter Smart Bombs + two ASRAAM.

Two ALOSNW (WE177B replacement) + two ASRAAM.

Two Special Elements (EMPW) + two ASRAAM.

10. Signature requirements given in the AIAA specifications are that the front

aspect RCS illuminated by a 1 – 10 GHz GCI, acquisition, and tracking radar

should be less than 0.05 m2. This would be demonstrated by SIGMA full scale

model, as for the F-35 described in appendices B.

11. The flyaway cost for 550 aircraft purchase must not exceed $75M (year 2001

dollars).

12. The IOC for the first operational squadron is 2020.



8

Figure 5: - AIA Baseline sizing mission. (This is based on reference 1).

13. The aircraft must be capable of using the following fuels JP-5 (RAF) (JP-5

Mil – spec density of 6.82lb/US gal equal to 51.10lbs/ft3) / JP-8 (USAF) (JP-8

Mil - spec density of 6.80lb/US gal equal to 50.86lbs/ft3) / and JP-4 (Special

fuel) (JP-4 Mil – spec density is 6.55 lbs/US gal equal to 49lbs/ft3) without

reducing range or mission effectiveness.

14. Maximum use is to be made of off platform sensor inputs for mobile target

acquisition (e.g. SCUD‟s, SS-20 IRBM‟s, and SS-25 ICBM‟s).

15. A UACV variant of the FB-24 designated A-24 of common core airframe was

required for the mixed fleet concept and this airframe is covered in this

conceptual design study the detailed systems study will form a subsequent

research project*.

*Covered in full in proposal submission document RFPS – 022007 which is not

part of the Cranfield MSc course.



9

Table 1:- Mission breakdown for the Advanced Interdiction Aircraft.

Segment. Description. Height. Speed. Distance/Duration.

1 - 2 Warm-up, taxi

and take-off. Sea – level.

NATO 8000ft icy.

2 - 3

Climb to best

super – cruise

altitude.

3 - 4 Cruise to conflict

area. BCA. BCM 522nm

4 - 5 Climb to

50,000ft.

5 - 6 Dash to target. 45,000ft Mach 1.6 378nm

6 - 7 Turn and

weapon release. 45,000ft Mach 1.6 180 degrees.

7 - 8 Dash out. 45,000ft Mach 1.6 378nm

8 - 9 Descend to super

- cruise altitude.

9 - 10 Cruise return. BCA. BCM 522nm

10 - 11 Descend to

airbase.

11 - 12 Land (with

reserve fuel*). NATO 8000ft, icy.

*Diversion and hold at sea level 30 minutes fuel at economical flight conditions.

(Reference 1(a): - Aircraft Design Projects for Engineering Students: by Jenkinson L.

R. and Marchman J. F. Published by AIAA Education Series in 2003. ISBN 1-56347-

619-3. Pages: - 208-209, and 1(b) USAFA course FSAV study included in

Aerodynamic Version 3.0 Software: Published by the USAFA and released by AIAA

Education Series in 2004: ISBN 1-56347-689-4.)

The A-24 UCAV used the same sizing mission defied above and was considered for

strikes on enemy chemical and biological research facilities with non-nuclear

weapons only, (the arming of UCAV‟s with tactical or strategic nuclear weapons

being politically unacceptable) this variant would be autonomous after control release

from the manned FB-24 (each FB-24 would support four A-24‟s) outside the key treat

zone. The FB-24 manned aircraft would control in-flight refuelling of A-24‟s when

required, conduct systems health checks, and mission management acting as fleet

commander in all mixed fleet operations. As a single strike component the FB-24

would conduct all nuclear strike missions and would be the only nuclear equipped

counterforce component of a mixed fleet operation.



10

2.1 Current F-35 family of aircraft.

The current F-35 Joint Strike Fighter family consist of three aircraft variants for three

United States military air arms and the United Kingdom armed forces, which have

common basic configurations and use cousin parts in most cases, are outlined below

in terns of lead dimensions: performance: and weight.

F-35A (CTOL Variant): - This is the USAF variant with 9g manoeuvring

capability and replaces the F-16 and A-10 (current orders are for 1763 aircraft).

Dimensions are: - Wing span 35 ft: length 51.1 ft: and wing area 460 ft2. Empty

weight is: - 27,395 lbs. Internal fuel capacity is: - 18,498 lbs. Combat radius is: -

greater than 590 nautical miles.

F-35B (STOVL Variant): - This is the USMC / U.K. RN & RAF variant to

replace the Harrier AV-8B / Mk - 9 and the F/A - 18 Hornet C/D aircraft (current

orders are for 759 aircraft). Dimensions are: - Wing span 35 ft: length 51.1 ft: and

wing area 460 ft2. Empty weight is: - 30,697 lbs. Internal fuel capacity is: - 13,326

lbs. Combat radius is: - greater than 450 nautical miles.

F-35C (CV Variant): - This is the US Navy aircraft carrier variant which will

replace the F/A - 18 Hornet C/D and the F - 14D Tomcat (current orders are for 480

aircraft). Dimensions are: - Wing span 43 ft: Length 51.4 ft: Wing area 620 ft2.

Empty weight is: - 30,618 lbs. Internal fuel capacity is: - 19,100 lbs. Combat

radius is: - greater than 600 nautical miles. Maximum payload all internal and

external stores stations used: - 21,400lbs.

Currently the F-35 family are to be fitted with either the 43,000lb maximum thrust

Pratt & Whitney F-135 or the GE Aircraft Engines / Roll Royce F-136 engines, which

will be sourced in alternate year USA Department of Defence equipment purchases.

The F-35 family fuselage OML is driven primarily by the need to house the STOVL

variants lift fan / drive shaft combination, and the two 2,000lb JDAM and two AIM-

120 missiles of the CTOL variant requiring a large cross sectional area, additionally

the common fuselage length is driven by deck spot requirements of both the STOVL

and CV variants.

The new FB-24 will not have the same commonality drivers because this airframe is

driven by the requirements of increased range and the ability to super cruise, however

it will use the F-35 systems and many common features developed on the F-35. This

type of reconfiguration follows the precedent set by the F/A-18E and F variants

which although larger than the original F/A-18 family share common baseline

configuration.

For this aircraft study the current F-135 and F-136 Pratt and Whitney, and GE

respective engines of 43,000lbs maximum wet thrust were the initial power plants of

choice which giving commonality with the current F-35 family, although an enhanced

version of the YF-120 VCE was to be considered with a military dry thrust of

30,000lbs if the former engines proved to be unsuitable.



11

2.2 Threat analysis.

Although the threat from the former Soviet Union has receded, most potential

aggressor nations have purchased, or inherited similar advanced air-to-air and surface-

to-air threat capabilities. Former Soviet Union fighter missile and airborne early

warning aircraft have been and continue to be sold to threat nations and these will be

assessed below.

The use of lookdown shoot down radars on fighters like the MiG-31 Foxhound

combined with Airborne Warning and Control System (AWACS) aircraft like the

Ilyushin / Beriev A-50 Mainstay, and the high performance of man portable surface to

air missiles like the Stinger, make the terrain masking attacks of the 1970‟s highly

questionable, and direct low level passes for airfield attack, as conducted by RAF

Tornado‟s in the first Gulf War suicidal.

The reaction time of defensive SAM systems, and air defence fighters is reduced if

the aircraft is stealthy, and the significant reduction of an aircrafts radar, infra-red,

visual and acoustic signatures, can greatly enhance its survival, resulting in grater

weapons system effectiveness. Although radar and infra-red stealth are the most

obvious signature features, other aspects of low observability (i.e. acoustic and visual)

cannot be ignored. The lessons learnt by the US air arms during the Vietnam War

emphasised the need for signature control, especially in the areas of Radar Cross

Section (RCS) and Infra – Red (IR) which will be covered below as well as visual and

acoustic signature reduction in general terms.

2.2.1 Air threats:-

For this study the primary air to air threats are anticipated to be MiG-29C‟s: MiG-

31M‟s: and control aircraft of the A-50 Mainstay type because in the future Global

Strike ConOps force in which this aircraft is intended to operate Eurofighter Typhoon,

and F/A-22 Raptor air dominance fighters will be responsible for tying down the Su-

27C / P, and Su-30A enemy air assets, when encountered.

Also with the latter aircraft being the most expensive Russian export platforms

nations capable of procuring them are most likely to employ them for free ranging

attack fighters rather than homeland defence tied combat air patrols. Each of the

probable threat aircraft capabilities are detailed in Appendices A, to the extent to

which published data is available.

The MiG-29 M shown in figure 6 is a major air threat to the AIA and from the list of

end users above the five nations highlighted in dark red are all potential near term

threat nations. Also this aircraft has been used against NATO forces in the Balkan‟s

war of the 1990‟s.



12

The MAPO MiG-29M: air defence fighter.

Figure 6:- The MiG-29 M, tactical air defence fighter which is considered to be

one of the best of the fourth generation fighters and is used by: - Belarus:

Bulgaria: Cuba: Czech Republic: Germany: Hungary: India: Iran: Kazakhstan:

Malaysia: Moldova: North Korea: Poland: Romania: Russia: Serbia: Slovakia:

Syria: Turkmenistan: Ukraine: USA(combat training): Uzbekistan: Yemen.

(Ref:-8 and 9).

The MAPO MiG-31B: interceptor fighter.

Figure 7:- The MiG-31B currently Russia‟s principal interceptor fighter as seen

at le Bourget, Paris air show and currently variants are to be offered for sale to

Syria: Iran: North Korea: and Yemen (Ref: - 8 and 9).



13

The MiG-31B figure 7 is capable of intercepting airborne targets in all weather

conditions, day or night, in continuous or intermittent radar coverage of Ground

Control Interception (GCI) sites, and is unaffected by the target aircrafts use of

electronic counter measures (ECM) or evasive manoeuvring. For operations outside

continuous GCI radar cover the MiG-31B can use its APD-518 (Apparatura Peredachi

Dannykh / Data Transfer Equipment) data – link system, both for reception of

guidance commands and for the distribution of target information (the air situation)

between aircraft in a formation. This would usually be a flight leader and three

wingmen. The MiG-31B combat radius is limited to a supersonic intercept radius of

389nm due to engine lubrication, and crew endurance, however at subsonic speeds

this is increased to 648nm, without overload tanks and up to 756nm when these tanks

are used.

Ilyushin / Beriev A-50 Mainstay AWAC

Figure 8:- The weathered surface of the rotating antenna of the Shmel

(Bumblebee) radar system is in contrast to the clean lines of this A-50, note the

fin – top fairing is for the Mnk (Poppy) missile approach warning system. Used

by: - Russia: India: and offered to Syria: Iran: North Korea: PRC: (Ref:-9)

The third and possibly the most important asset to the potential enemy nations is the

Ilyushin / Beriev A-50 Mainstay long range airborne warning and control system

(AWACS), shown in figure 8, (developed jointly, as the designation indicates, by the

Ilyushin Design Bureau and the Beriev Aviation Scientific and Technical Complex at

Taganrog on the sea of Azov), as a airborne interception control platform this aircraft

is a priority target for the air launched low observable anti – radiation weapon.

2.2.2 Surface to air threats: -

Russia is selling its latest and best systems on the world market and the two mobile

weapons of choice purchased by the threat nations are outlined below, neither of

theses can easily be destroyed by the X-45 or the X-47 UCAV‟s or indeed sea

launched TOMAHAWK cruise missiles.



14

Primary surface to air threats to the FB-24 will come from the mobile S-300V SAM

system, and the highly mobile Buk-M1 (SA-11) replacement for the Kub (SA-6)

SAM that shot down Captain Scott O‟Grady‟s F-16 over Bosnia.

The S-300V (SA-12): - Gladiator SAM system.

Figure 9: - The S-300V (SA-12) missile system which can be deployed or

dismantled and moved in five minutes. Developed in the 1980‟s the full S-300V

system has been exported to the following nations: - Syria: Iran: North Korea:

PRC: (Reference2: -Pages 87-88, F-22 Raptor: by Sweetman B: Published by MBI

Publishing company USA 1998)

The S-300V (SA-12A Gladiator) system shown in figure 9 has two missiles

(developed from the anti-ballistic missile SA-12B Giant SAM), one large and one

small, the smaller one has a peak velocity of Mach 6 and can destroy targets evading

at 8g through clutter and ECM / decoy systems over an effective range of 30nm at

altitudes between 2,000ft and 60,000ft, the second larger missile attains a peak

velocity of Mach 8 and through advanced terminal aerodynamics can destroy targets

manoeuvring at 12g through ECM / decoy systems at altitudes between 12,000ft and

80,000ft. Targeting information can be obtained from the main Almaz NPO family of

sensors or off bored from the A-50 Mainstay or MiG-31 Foxhound airborne

platforms, and are capable of intercepting HARM missiles launched against them.

In addition the Almaz S-400 Triumf (SA-20 Gargoyle) family which are improved

faster derivatives of the S-300V‟s are now available on the world market after

completing field trials. (Reference 3:- pages 284-285, Iron Hand Smashing the

Enemy’s Air Defences: by Thornborough. M. A. and Mormillo. B. F.: Published by

Patrick Stephens Limited an imprint of Haynes Publishing UK 2002)



15

The Buk-M1 (SA-11):- SAM system.

The Buk-M1 is the replacement for the Kub (SA-6) and is in full scale production for

the export market, with a single system which is mounted on 11 vehicles a defender

has 36 missiles ready to launch at any one time. The missiles have a reach of some 45

to 105nm, depending on model similar to that of the SA-10 Grumble, with a closing

speed of up to 14,770ft/sec. Also like the SA-10 this weapon system relies on radar

guidance coming from associated F-Band Continuous Wave pulse – Doppler Clam

Shell, tower - mounted Big Bird or 3-D Tombstone long – range surveillance / EW

radars, with I/J-Band Flap Lid phased array radar used for target tracking. As with the

S-300V and S-400 this system can be linked to the A-50 so that target information can

be obtained without the risk of the system being exposed to attack from anti –

radiation missiles.

Short – range Man portable SAM‟s.

The man portable SAM‟s MANPADS of the SA-14/-15/-16 have imaging infra-red

and ultraviolet seekers operating at both ends of the visible spectrum, and the SA-16

Gimlet simply ignores flares altogether, and over 1 million have been sold to date

most to threat nations. The best counter to these is to fly above 20,000ft to avoid them

and anti aircraft artillery (triple –A) fire, and rely on older precision guided munitions

PGM‟s or newer JDAM‟s and J-series weapons in a fast pass lob well away from the

target area. But such tactics place non-stealthy aircraft within the prospective shooting

range of the SA-11 SA-10, and S-300V and S-400 SAM‟s detailed above. (Ref 3)

Against these new threats the time-honoured defence tricks such as terrain masking

and defensive high – g breaks simply would not work against S-300 class SAM‟s

even if the pilot was G-LOC immune: these weapons use gas-dynamic control

systems for super-manoeuvrability during the closing stages of interception, which

allow a 20g acceleration in under 0.025 seconds in response to such tactics, as well as

the ability to engage targets down to treetop level (assuming adequate line-of-sight for

a tower – mounted tracking radar). Commenting on this surface to air threat

environment one pilot put it, “You have to hide from them until you can kill them”.

New fourth generation stealth fighters such as the F-35 Joint Strike Fighter and F/A-

22 Raptor, according to published texts relying on golf ball and marble Radar Cross

Section‟s (RCS) respectively, aim to do precisely that – hide, not evade, and right

over the enemy‟s noses. General John P. Jumper stated “High – altitude attacks

at up-to 40,000ft would help them avoid the operational envelope of infra-red

seeking SAM‟s and triple A, and speeds of Mach 1.3 – to – Mach 1.5 would

reduce the effective envelope of the large radar guided SAM‟s like the S-300V‟s,

S-400‟s, and SA-11‟s” (Ref 2 and 3), which is the most important factor in

determining the mission profile of the FB-24 AIA, and hence the aircraft itself.

2.3.3 Stealth requirements.

This is the ability of an aircraft to attack its target with the maximum amount of

surprise by denying long - range detection and is vital to the FB-24‟s interdiction and

strike missions. This is achieved by reducing the aircrafts visual, radar and infrared

signal strengths of which a detailed treatment of these is given in Appendices B, and

an overview is presented here.



16

1 Visual: - Elimination of smoke trail with and without afterburner. Camouflage the

aircraft by painting it in colours predominating in its mission environment, and reduce

the size, number and visibility of identification markings.

2 Radar: - Reduce radar cross – section (RCS) by: avoiding surfaces at right angles to

each other (to limit the number of corner reflections), designing for a minimum of

radar spikes (by minimising the number of airframe angles), shielding the engine

compressor face, using radar absorbent materials, carrying stores internally, treating

the cockpit canopy and radar cavity, minimising energy emissions from the aircraft‟s

own sensors, providing comprehensive electronic countermeasures and decoys.

3 Infra – red: - Less use of afterburner, with its enormous IR signal, since the very

short afterburners now employed have increased the angular detection range. Shield

and / or cool engine exhaust, and where appropriate use tuned decoys which match the

emission spectra of the aircraft.

As is shown below in the rest of this thesis these requirements have a profound effect

on the design of the FB-24 both externally and internally, as well as systems selection.

3.0 FB-24 and A-24 configuration studies and design selection.

3.1 Initial Conceptual Design Studies.

In order to meet the Advanced Interdiction Aircrafts basic requirements three options

were considered (as per Reference4:- Page-7: Aircraft Conceptual Design Synthesis:

by Howe. D: Published by Professional Engineering Publishing Ltd 2000) and are

detailed below: -

1. Adaptation or a special light version of the existing F-35C by removing all

carrier born equipment and structural requirements from the airframe and

adopting a new larger wing, with a modest forward fuselage extension for the

second crew member, as a low cost, low risk option retaining a high degree of

commonality with the F-35 family. However this conservative approach would

not meet the supercruise capability requirements as the: - fuselage finesse ratio:

wing plan form and sweep angle: and greater wetted area, would induce more

drag and would give a similar performance to the original F-35C which is not

supercruise capable. This in turn would reduce the effect of increasing the fuel

volume on range as large portions of the mission would require afterburner use

to meet the required Mach number for ingress and egress of the target zone.

2. A major modification or direct development of an existing type this option

involved a major redesign of the existing airframe, consisting of: - extensive

fuselage extension to increase the finesse ratio: wing sweep and planform

alignment changes: empennage changes: to reduce drag, and the removal of all

non – land based strike equipment: combined with a reduction in airframe

substructure component weight reflecting the more benign operating

environment.



17

This was a much more radical approach which was more expensive and sharply

reduced commonality with the rest of the F-35 family in major airframe

components, although internally 40% of the substructure would be cousin parts

with the CTOL variant and the undercarriage would be identical to the F-35C.

The level of systems commonality with the CTOL variant would be

approximately 80% (are estimated to be 45% of total platform costs reference 5

page 21 Fundamentals of Fighter Design 1st Edition: Whitworth. R:

Published:- Airlife 2003), with specialised systems for the UCAV command

and nuclear strike roles. This approach had a higher probability of meeting the

Advanced Interdiction Aircraft requirements than option 1, with a high degree

of commonality in the expensive system components of the platform e.g.

offensive and defensive avionics, EHA‟s, and fibre optic cable data links, as

well as a degree of structural component commonality. However this was a

considerably more expensive and higher risk option.

3. A completely new design this option was to produce a completely new aircraft

using two YF-120 Variable Cycle Engines in a much larger airframe optimised

specifically for the AIA and FSAV missions, incorporating smart structures

and new materials and manufacturing techniques, as well as specific

missionized systems. This would have no commonality with the F-35 family

airframe sub structure, and only some systems commonality. This option could

defiantly meet the AIA and FSAV requirements being specifically designed to

do so but would be too expensive for the production run envisaged and the cost

target would result operational support and logistics problems for F-35 owners

as a new aircraft type. The inherent risks in a completely new aircraft would

also very high and the development cycle would be too long based recent on

legacy projects like the Eurofighter Typhoon, FA-22A, in a changing military

environment. Although for the industry as a whole and national security it is

prudent to maintain an indigenous capability by embarking on a completely

new aircraft design and manufacture using the latest advances in technology

every twenty to thirty years.

On balance the considered decision was taken to pursue option 2 major

modifications to the existing F-35C to develop a platform capable of meeting the

detailed AIA and FSAV requirements within the cost limits and within the timescales,

required for this platform.

To map the customer requirements and determine the level of importance of

individual elements of the overall requirements to the customer and the engineering

solutions available the Quality Function Deployment method was used for this

project. This methodology incorporates the following: -

1. Language understood by all participants:

2. Cross functional cooperation:

3. Focused technology development:

4. Cost / benefit analysis.



18

The key benefits of QFD are as follows: -

1. Reduction in engineering change:

2. Shorter design cycles:

3. Lower start up costs:

4. Systematic documentation of engineering knowledge:

5. Competitive pricing:

6. A more satisfied customer.

AERO 481 QFD Process – House of Quality

1. Customer needs (whats)

2. Customer priorities

3. Technical solutions (hows)

4. Relationship matrix

5. Technical priorities

6. Target values

7. Correlation matrix

Design Feature

Correlation

Matrix

Customer

Priorities:

Weight Rank

Design

Features

Cu

stom

er N

eed

s

Design Feature

Priorities

Target Values

Figure 10: - The QFD Process illustrating the methodology of House of Quality

construction. (Slide source Reference 1: - USAF Academy lecture AERO 481,

incorporated within AeroDYNAMIC Version 3 software: by: - Brandt. S. A,

Stiles. R. J, Bertin. J. J, Whitford. R.: AIAA Education Series: Pub 2005: ISBN 1-

56347-689-4).

For this project a House of Quality was produced to determine the customer priorities

against each customer need and technical solution and is shown in figure 11 below.



19

Po

siti

ve

Syn

ergy

Inte

rfer

ence

neg

ativ

e

Str

on

g

Po

siti

ve

Ho

us

e o

f Q

ua

lity

De

sig

n F

ea

ture

sL

igh

tH

igh

AR

Lo

w A

RA

dv

an

ce

dL

ow

SF

CL

ow

RC

SIn

tern

al

Cu

sto

me

r N

ee

ds

C

usto

mer

Needs

We

igh

tW

ing

sW

ing

sE

ng

ine

Lo

w I

Rw

eap

on

sP

rio

ritie

sR

ank

Lo

w C

os

t9

36

02

20

0.0

80

1L

on

g R

an

ge

18

00

nm

08

29

93

90

.08

02

Ea

sy t

o S

erv

ice/M

ain

tain

50

36

33

20

.06

03

Su

rviv

ab

le6

00

64

99

0.0

80

4O

pe

rati

on

fro

m i

cy 8

00

0ft

ru

nw

ay

60

92

99

90

.06

55

Da

ma

ge

to

lera

nt

str

uc

ture

90

84

08

70

.05

06

All

we

ath

er

cap

ab

ilit

y8

15

55

08

0.0

80

7S

tru

ctu

ral li

mit

lo

ad

s 5

0%

fu

el +

7g

to

-3

g5

08

44

56

0.0

50

8M

axim

um

sp

ee

d M

ach

1.6

20

99

86

80

.05

09

Sta

bil

ity m

ee

ts M

IL-F

-878

5B

36

50

20

60

.04

010

Ac

co

mm

od

ate

tw

o c

rew

40

04

55

20

.06

311

Op

era

tio

na

l c

om

ba

t c

eilin

g 4

5,0

00

ft7

75

88

68

0.0

62

12

Op

era

tio

na

l c

om

ba

t ra

diu

s =

900

nm

78

59

95

80

.08

013

Ma

x i

ns

tan

tan

eo

us

tu

rn r

ate

8.0

de

g/s

ec

80

86

32

50

.04

014

SE

P (

1g

) d

ry t

hru

st

1.6

M a

t 4

5,0

00

ft=

0ft

/s8

08

76

05

0.0

30

15

SE

P (

1g

) w

et

thru

st

1.6

M a

t 4

5,0

00

ft=

200

ft/s

80

87

50

50

.03

016

SE

P (

2g

) w

et

thru

st1

.6 a

t 4

5,0

00

ft=

0ft

80

87

52

50

.03

017

Co

mm

on

ality

wit

h F

-35 f

am

ily

60

98

42

20

.03

018

D

esig

n F

eatu

re P

riorities

5.9

16

2.2

74

5.2

75

5.5

18

5.2

96

4.1

22

6.0

47

1.0

00

Ch

ecksu

m

T

arg

et valu

es

71,0

18

lbs

6.0

02

.98

2010

0.7

lb/h

/lb

f-2

0d

b5

,65

4lb

s

Fig

ure

11:

- F

irst

tie

r Q

FD

Hou

se o

f q

uali

ty c

hart

foe

the

Ad

van

ced

In

terd

icti

on

Air

cra

ft, p

rod

uce

d i

n A

erod

yn

am

ic V

3.0

.



20

3.1.1 FB – 24 and F-35C common features.

The following common features from F-35C were incorporated in to the

configurations modelled are described below:

Primarily the FB – 24 will be a single engine aircraft which will carry all

weapons and fuel internally, it will use the same undercarriage, avionics, and

systems as the current F–35C, but the developed aircraft will carry an internal

M61A1 20 mm Cannon for self protection. Also for all FB-24 missions will

use the ASRAAM as the self protection against air threats instead of the F-35C

AIM-120C advanced medium range air to air missile, because the FB-24 is not

intended to instigate air to air combat relying primarily on stealth for

protection.

The Diverterless Supersonic Inlet DSI developed for the F-35 will be carried

over to the FB-24 AIA, because the DSI effectively eliminates one of the major

RCS contributors in the frontal cone of modern high speed combat aircraft

namely the diverter system as outlined in Appendices C.

The use of foam filled fuel tanks employed on many current fighters including

the F/A-18E/F and proposed for the F-35 will also be employed for the FB-24

AIA rather than basic self – sealing tanks called up in the requirements reduces

the hydrodynamic ram effects caused by the impact and penetration of a

missile fragment through the fuel tanks

The GE F-136 / LOAN combination which is a low cost, light weight means of

achieving signature control in both RCS and IR spectrums, while providing

significant improvements in reliability, maintainability, and supportability,

compared to previous production nozzles. Reduction in RCS and advanced

material technologies allow axisymmetric nozzles to achieve signature levels

previously possible only with two dimensional (2-D) vectoring nozzles.

Advantages of the axisymmetric designs, attributable to inherent structural

simplicity and efficiency, have been employed to achieve substantial weight

and cost reductions compared to 2-D designs. In addition to geometrical

shaping and special materials for signature control, the LOAN also

incorporates an ejector that enhances nozzle cooling.

Durability has been significantly improved, and maintenance – friendlily features

reduce the time required to change components by as much as 80%. These features

strongly influence the selection of this combination for the FB-24 AIA. (Reference 6:-

Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by Harkins H:

Published by Centurion Publishing UK 2004)

The internal weapons carried for primary use will be two of either of the

following: - GBU-31 JDAM PIP Mk – 84 warheads: AGM-154 JSOW glide

bomb: Saber 15-Kiloton ALOSNW (not shown but of similar dimensions to

JSOW): in combination with two AIM-132 ASRAAM missiles requiring a rail

launch trapeze, in place of the AIM-120C AMRAAM‟s of the F-35C.



21

Figure 12: - Internal weapons carriage options of existing F-35 family retained in part

for the AIAA FSAV concepts. Source: - Reference 2:- Code One Lockheed Martin

Aeronautics publication.

This design study will use the government furnished equipment GFE listed below in

table 3 for weight and airframe OML sizing. Nine configurations were studied from

which a final basepoint configuration was selected the evaluation of these

configurations and how they met the merit measurement criteria is detailed below.

3.2 FB – 24 and A-24 Configuration concepts.

The key design drivers identified from the QFD house of quality analysis (figure 11)

influencing the OML of the new build (NB) AIA concepts considered under the

major modification approach, and the methods of their resolution are detailed

below: -

1. Long range at high speed, to be achieved through supercruise performance:

2. Stealth as defined in section 2 and detailed in appendices B, to be achieved by

planform alignment, and engine emission shielding:

3. Payload capability, to be achieved by stretched weapons bays:

4. Additional crew position, to be achieved by fuselage stretch:

5. Lower cost than a completely new design, to be achieved by retaining a degree

of commonality in systems and airframe with both the F-35 family, and the

F/A-22A.

Design drivers 1 through 5 have a direct impact on the OML configuration of the FB-

24 and are covered below in the initial configuration stage, and some elements of

design driver 5 are incorporated in resolving the first 4 design drivers.



22

Proposed resolution of Design driver 1:- At supersonic speeds aircraft drag is

composed of: - (a) Skin friction drag: (b) Wave drag due to thickness (or volume drag

also known as zero lift wave drag): and (c) Drag due to lift (a combination of vortex

drag and drag due to lift). The FB–24 configurations attempted to address the wave

drag of the current F-35C configuration by increasing wing sweep angle which raises

Mcrit and reduces wave drag, (but degrades CL and CLmax, and increases induced drag)

and increasing the ratio of the aircrafts maximum cross – sectional area to overall

length which has a larger effect on the configurations supersonic wave drag than dose

the wing sweep or the smoothness parameter EWD also the wing span and sweep angle

was engineered to contain the wing span inside the shock wave cone generated by the

aircraft‟s nose, further reducing the wave drag because Mach number inside the cone

is lower than the freestream Mach number (M), and shock waves are weaker than

they would be if the wing were exposed to M. The fuselage had area rule

configuration to reduce drag at transonic speeds and increase transonic acceleration,

with a blended wing, fuselage interface.

The wing design for Advanced Interdiction Aircraft gives rise to multiple design

points as with all tactical military aircraft (Reference 7: Bertin. J. J: Aerodynamics for

Engineers 4th

Edition: USAFA: Published by Prentice Hall 2002) and this multiple

design point requirement was a major driver, with the aerodynamic requirements for

each point often in conflict with each other. For example the need for rapid

acceleration to supersonic flight and effective supersonic cruise requires a thin wing

section with relatively high sweep and with camber that is designed to trim out the

moments resulting from the aft aerodynamic centre movement at supersonic flight.

However, these requirements conflict with those of efficient transonic manoeuvre

which is better performed with thicker wing sections designed with a camber for high

CL operation and a high aspect ratio planform to provide a good transonic drag polar.

The final design is therefore a compromise solution of a variable camber wing was

considered in preference to variable sweep wings, as on Tornado and F-111 which

have prohibitive volume, weight, complexity, and stealth issues and would be

completely inappropriate for the AIA. A typical performance spectrum for tactical

fighters corresponding to typical mission requirements is shown in figure 13 which

presents a map of lift coefficient versus Mach number. The low Mach end of the

spectrum throughout the CL range is typical of takeoff and landing for the

configuration. The subsonic cruise and supersonic cruise portions are noted in the

moderate lift range. Acceleration to high supersonic speed occurs at low lift

coefficients Sustained manoeuvre takes place in the CL range of less than one for most

fighter configurations, and above this lift coefficient, the aircraft is in the

instantaneous manoeuvre regime. Drag rise occurs depending on the wing geometry,

in the range of Mach 0.8 to 1.2. The particular flow conditions that correspond to the

flow map are shown in figure 14.

At the cruise and acceleration points the primary consideration is attached flow, and

the design objective is to maintain attached flow for maximum efficiency. At the

higher CL values corresponding to instantaneous manoeuvre, separated flow becomes

the dominant feature. Current designs take advantage of the separated flow by forming

vortex flows in this range. Intermediate CL values corresponding to sustained

manoeuvre are usually a mixture of separated and attached flows.



23

Figure 13:- Typical performance map for a tactical fighter aircraft, (Source

Reference 7 Page 514).

Figure 14:- General flow regimes encountered for tactical fighters, (Source

Reference 7 Page 515).

Consequently, if the aircraft is designed with camber to minimise separation in the

manoeuvre regime, the configuration will have camber drag and may have camber

drag and may have lower surface separation, which increases drag at the low CL values

needed for acceleration.



24

Therefore the overall wing design for the AIA needed to be a compromise to achieve

optimum flow efficiency through all of the design points of the proposed mission, and

the design was influenced to the greatest extent by the constraints imposed by

supersonic and high subsonic cruise.

For the wing design the ideas of Bradley and Bertin (Reference 7 pages 516 - 517)

were employed as follows, Bradley states in referring to the combination of variable

leading edge flaps with strake configurations: - “We should mention a new concept of

controlled vortex flow for designing military aircraft having emphasis on supersonic

configurations. Wing planforms for supersonic cruise have higher leading edge sweep

and generally lower aspect ratios: these planforms develop vortex flows at relatively

low angles of attack. As a result, the transonic drag characteristics are lacking in the

manoeuvre regime since drag polars generally reflect very little leading edge suction

recovery. Recently, wings of this type have been designed to take advantage of

separated vortex flows rather than to try to maintain an attached flow to higher CL

values.” Small leading edge closeouts incorporated in the YF-22, and both X-35

demonstrator aircraft and the fore body chine of the YF-23 had the potential to create a

vortex flow field, although not anywhere as large as those of the F-16.

Bertin in considering variable – twist, variable – camber wings states: - “Survivability

and mission effectiveness of a supersonic – cruise military aircraft require relatively

high lift / drag ratios while retaining adequate manoeuvrability. The performance of a

moderate – aspect ratio, thin swept wing is significantly degraded at high lift

coefficients at high subsonic Mach numbers because of shock – induced boundary –

layer separation and, at higher angles of attack, because of leading – edge separation

and wing stall. The resulting degradation in handling qualities significantly reduces the

combat effectiveness of such airplanes.” The techniques used to counter leading –

edge stall, are as follows: - leading edge flaps, slats, and boundary – layer control by

suction or by blowing (used on the Blackburn Buccaneer, and Lockheed F-104), which

have been effectively employed in combination with trailing edge flaps for increasing

the maximum usable lift coefficient for both the low – speed landing, and high

subsonic phases of a tactical strike aircraft mission. For supersonic drag and stealth

considerations, boundary layer control was discounted and full span constant hinge

line leading edge flaps were employed in combination with trailing edge flaps for

landing and high subsonic phases of the AIA mission this is common to the F/A-22A

and the F-35 family. Bertin recommends low – thickness – ratio wings incorporating

variable camber and twist for high performance fighters with fixed wing planform,

because the camber can be reduced or reflexed for the supersonic (dash to target)

phase of the mission and increased to provide the high lift coefficients required for

transonic and subsonic manoeuvrability, ideally a smart structure mission adaptive

wing would have been the logical selection, however the research and development is

still at a relatively fundamental stage and this technology was not deemed mature

enough for the FB-24. A test program was conducted by NASA in 1977 (Reference8: -

Ferris, J.C.: “Wind-Tunnel Investigations of Variable Camber and Twist Wing”:

TND-8457: NASA: Aug 1977) to determine the effect of variable camber and variable

twist on the aerodynamic characteristics of a low – thickness – ratio wing. The basic

wing was planar with a NACA 65A005 airfoil at the root and a NACA 65A004 airfoil

at the tip (in effect there was no camber or twist in the basic wing). The section

camber was varied using four leading - edge segments and four trailing - edge

segments, all with span wise, individual hinge lines as shown in figure 15.



25

Figure 15:- The demonstration wing configuration used to show the effects

variable twist and camber in the 1977 NASA experiments (Source Reference 7

Page 518).

Variable twist was obtained by having greater sweep angels in both the leading – edge

and trailing – edge segments than in the basic planforms leading and trailing edges, as

shown in figure 15. All leading – edge segments were parallel to each other as were all

of the trailing edge segments. Camber and twist could be applied to the wing as shown

in figure 15. The test program (Reference 8) demonstrated that deploying the trailing

edge segments near the wing root crated a cambered section with an effective chord of

increased incidence, where as deploying the leading edge segments near the wing tip

creates a cambered section whose local incidence is decreased. Hence the modified

wing could have an effective twist of approximately 80 of washout. The program

further demonstrated that the use of leading edge camber lowers the drag substantially

up to a lift coefficient of 0.4, and increases the lift / drag ratio over a Mach number

range 0.6 to 0.9, and at lift coefficients above 0.5, the combination of twist and camber

achieved combining both the leading - edge and trailing - edge segments was effective

in reducing drag. Trailing - edge camber was also shown to cause very large

increments in CL with substantial negative shifts in the pitching moment coefficients.

Examples of the data obtained from the NASA research is shown in figure 16, and it

can be seen that the maximum lift / drag ratio for this particular configuration at M =

0.80 is 18 and occurs when CL = 0.4.



26

Figure 16:- Effect of twist and camber on the longitudinal aerodynamic

characteristics determined from the NASA test program (Reference 8) chart (a) is

lift-to-drag ratio and drag polar, and chart (b) is pitching moment and lift

coefficient.

The airfoil section selected for the AIA was NACA - 0006 which was a symmetric

airfoil with a maximum thickness of 6% of the chord with a sharp leading - edge.

Which although thicker than the F/A-22A airfoil section which has a customised

airfoil thickness of 3.8% optimised for transonic operation, has similar root thickness /

chord ratio of 0.06 compared to the formers 0.592 and a tip thickness / chord ratio of

0.429 (Reference 9:- Miller. J.: Lockheed Martin F/A-22 Raptor: Stealth Fighter:

Published by Midland Publishing 2005) the thickness selected for the AIA wing was

within the transonic and supersonic efficiency thickness range of 5 – to – 8% required

for a multi – role strike / fighter airfoils as stated by S. Kern (Reference 10:- Kern. S.:

Evaluation of Turbulent Models for High – Lift Military Airfoil Flowfields: AIAA96-

0057: presented at the 34th

Aerospace Science Meeting, Reno, Nevada: Jan 1996):-

“Integration of stealth requirements typically dictates sharp leading edges and

transonic and supersonic efficiency dictates thin airfoils on the order of 5-8% chord”

initially the intention was to evaluate the effects of camber and twist variation on the

three wing planforms selected in section 2 with the NACA 0006 airfoil using the Flite

3D CFD software package however the BAE JSF IPT would not to support the authors

training for this activity and this is an area of future study and is now outside the scope

of this thesis. The degrees of twist and camber selected were based on F/A-22A values

from reference 9 and MDA / BAe JAST configuration 9B data (Reference 6:-

Hawkins. H.: Lockheed Martin Joint Strike Fighter (The Universal Fighter):

Published by Centurion Publishing 2004), and F-16XL from (Reference11 / 12:-

NASA Dryden Fact Sheets – F-16XL-1 Testbed Aircraft:

www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html).

As illustrated fig 15 As illustrated fig 15

http://www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html



27

To further reduce drag four different empennage configurations were investigated in

combination with the three wing planforms which were as follows: - the current F-35

and F/A-22A Four tail layout i.e. Vertical tails canted outward 250 to 27

0 from the

aircraft z, x, plane matching the slope angle of the fuselage side walls with fixed

torsion box and leading edges and attached rudders, and Horizontal tails which were

all moving with leading edge sweeps matching that of the wing, attached either by a

root spigot or integral hinge spar: or the YF-23 all moving Ruddervators canted out

500 from the aircraft z, x, plane matching a localised rear fuselage side wall slope, and

attached with a root spigot. A canard configuration considered without the vertical

tails observed in the Lockheed Martin JAST studies (V- vertical tails) and Typhoon /

Rafale etc (single vertical), similar to the X-36 was also considered. Finally a

completely tailless and canard less configuration with no empennage group at all as

observed on the X-45 and X-47 UCAV‟s.

The first conservative four tail empennage option was readily applicable to large

conventional missionized wing and the cranked arrow head MDA / BAe JAST wing

but not to the double delta X-16 wing and would have had a similar drag to

contribution to the former wing configuration as the F-35C empennage has to the F-

35C, and would be sized proportional to the wing.

The second more radical Ruddervator option was applicable to all three wing

configurations and offered a real drag reduction possibility, on a developed F-35C /

FB-24 configuration. Also this empennage configuration was flight proven in

supersonic, supercruise, and low speed flight regimes by the Northrop YF-23 during

the Advance Tactical Fighter demonstration and validation phase test program in 1991

and is shown in figure 17 below.

Figure 17:- Northrop McDonnell Douglas YF-23 showing 500 canting of the

ruddervators applied to all three of the FB-24 wing configuration. Source:-

Authors private collection.



28

The third canard configuration considered for the FB-24 AIA, was the short coupled

arrangement with the wing planforms considered, in this arrangement the foreplane

was to be located just ahead of and just above the wing like the X-36, Rafale, and the

original LM JAST configuration pre - F-35. Carefully positioning the canard and wing

relative to each other enables their combined lift effectiveness to exceed the sum of

their individual values. This configuration had the possibility of higher agility than the

first conventional configuration and potentially a lower drag contribution. However

the size (lager than on either the Rafale or JAST configurations, because of not having

any vertical tails) and location of the canards would be detrimental to the frontal RCS

(see appendices B) of the aircraft because of the increased number of reflecting

surfaces, every time the canards are actuated.

Although the fourth completely tailless option theoretically offered the lowest drag,

longitudinal stability was thought to be inadequate due to the short moment arm of the

elevators in all but the delta design, requiring a multi axial vectored nozzle and

complicating the wing flight control surfaces. The X-45 shown in figure 18 has a short

and flattened fuselage shape of 26 feet 5 inches, with the wings of 33 feet 8inches span

located at the rear where the elevators are most effective, where as the X-47 is a

diamond shaped flying wing and incorporates additional control surfaces to elevators

and flaps, buried in the wing top and bottom surfaces. Although only the X-45

employs a yaw thrust vectoring system, neither is supersonic, and there configurations

are totally different to any considered for the FB-24, the only similar configurations to

the FB-24 for which the tailless approach has been advocated i.e. the X-44 MANTA

shown in figure 19 and the FB-22 both use vectored thrust and have two engines to

effect pitch and roll control, therefore this empennage less configuration was not

considered further in this concept design study.

Figure 18:- The X-45 tailless UCAV although not readily apparent from this view

the aircraft has a short fuselage compared to its wing span with pitch, yaw, and

roll being controlled by a total of six trailing edge control surfaces. Source: -

USAFA AeroDYNAMIC Version 3.0 CD-ROM.



29

Figure 19:- X-44 MANTA F-22 applications project although not built or flown

wind tunnel models and CFD simulations demonstrated the requirement for

continuous vectored thrust augmented stability and to achieve this, a two engine

2-D vectored nozzle was necessary. Source: - Authors private collection.

Therefore in the final analysis only the conventional F-35 and F/A-22 and the YF-23

ruddervator empennage configurations were considered suitable for the FB-24 design

constraints and had the merit for further evaluation in this conceptual design study.

Proposed resolution of Design driver 2:- As stated in Lockheed Martin‟s Affordable

Stealth paper (Reference 13) and reinforced in references 4 through 7 and detailed in

appendices B, In order to develop a low RCS aircraft consideration must be given to

any part of the aircraft that a radar wave can reach. This impacts on the OML design

because the shape of the aircraft is the most critical factor to consider in the design

process for LO capability. A visual study of the F-35 and F/A-22A s reveals that all

hard edges i.e. the wing leading and trailing edges, control surfaces, vertical and

horizontal tail surfaces, intakes, have been aligned to a few common angles, so that the

radar returns from them all point in the same few directions. This is called planform

alignment and is carried over doors, external sensor apertures, inspection hatches, and

even formation and navigation lights, where this is not possible multiple chevrons are

used to break up the radar return. The overall result is a few relatively large but narrow

signature spikes that are difficult to detect and track.

The vertical tails and fuselage sides are tilted the side walls meeting the top surface in

a continuous chine running from the tip of the radome to the rear tip of the fuselage

this avoids a direct radar reflection from side on illumination. The entire surface is

blended smoothly at major component interfaces, enabling electrical surface currents

to flow over the aircrafts surface without interruption, in addition this also has the

effect of reducing drag. The effects of breaks for control surface interfaces are reduced

by radar absorbent seals both on the control surface and on the main plane trailing

edge. This planform alignment was employed in all of the study configurations

proposed in this thesis.



30

One of the largest remaining radar reflectors is the cockpit with potential radar returns

coming from the pilots and offensive air systems officers heads and helmets, seats, and

all of the various controls and displays in the cockpit. The most effective way to

reduce the cockpit signature is to prevent the radar energy from entering the cockpit. A

metallic coating is currently applied to all USAF, USN and USMC combat aircraft

which prevents the energy from entering the cockpit, eliminating this concern. The

cockpit frame however is still a potential area of concern and on both F/A-22, F-35 the

frame front follows the planform alignment principal, and in the former aircraft so

dose the frame rear. The cockpit selected for the FB-24 was a two place cockpit based

on the cancelled F-22B training version of the F/A-22A, unlike the F-35 which is

largely driven by the STOVL pilots vision requirements, this cockpit had high sills

enabling smoother blending into the fuselage of the two piece canopy which has an aft

swept bow frame to resist birdstrike this being required because of the increased

length over that of the F/A-22A‟s single piece canopy.

Another major contributor to the aircraft signature would come from the intakes, in the

FB-24 the same intake design features of the F-35 were employed namely the

diverterless supersonic intake, which in the JSF support program was demonstrated to

be suitable for flight in the Mach 2 flight regime (see appendices C for details), and

the characteristic forward sweep to the top of the intake to allow diverted air to spill

out over the local fuselage top surface. The intake ducts are bifurcated as in the F-35

resulting in complete engine face obscuration.

The rear quadrant signature reduction for the FB-24 was achieved by employing the

LOAN nozzle however because the detailed of the nozzle is secret a generic volume

cone of appropriate dimensions was used for all models. The intension was to shield

side view of the engine bay and nozzle with the empennage and this was successfully

accomplished for both four and two piece tail configurations. The engine and nozzle

was situated further aft in relation to the empennage group than is the case with the F-

35 where location is effected by STOVL commonalty requirements.

The final external feature is the Air Data System (which caused an extensive research

work on the worlds first stealth combat aircraft the F-117A), for which the FB-24

adopts the F/A-22A low observable pneumatic air data system (PADS), which

consisted of two small facetted fuselage mounted air data probes, one on either side of

the fuselage located aft of the radome, and four flush - mounted static ports (two on

each side of the fuselage) which were also located aft of the radome above and below

the forward fuselage chine, eliminating the multitude of probes seen on non stealthy

aircraft such as Tornado.

These OML features combined with the structural layout and materials features

covered in section 4 were deemed to resolve design driver 2 for the FB-24 conceptual

design study. (Reference 13: - Lockheed Martin’s Affordable Stealth: paper by Haisty.

B. S.: Published by Lockheed Martin Aeronautics Washington D.C.: November 15th

2000: for National Press Club).



31

Proposed resolution of Design driver 3:- The weapons bays were positioned so that

its centre of gravity was as close as practical to the whole aircraft centre of gravity. A

single weapons bay was ruled out for two main reasons:- firstly with a single engine

the duct from the bifurcated intakes ran through the aircrafts centre line therefore a

single weapons bay would have to be positioned below this and result in a dramatic

increase in fuselage cross section impacting adversely on the desired finesse ratio:

secondly the air turbulence generated when a single deep weapons bay was opened at

supersonic speeds, resulting in acoustic damage to the internal structure, and adverse

handling. Therefore two weapons bays one on each side of the fuselage, each with

inboard and outboard doors as on the F-35 family was considered to be the best

solution for the FB-24 however the length was increased from approximately 4.5m to

7.5m to accommodate both a GBU-31(v) 3/B (PIP) and a ASRAMM in tandem, in

each by, with the former being released by cold gas ejectors and the latter being

launched from the integrated outboard door launch rail. The rear bulkhead of each

weapons bay was sloped rearwards to improve airflow when the bays are open and to

give better nose down pitching characteristics should the aircraft force land in water

and the weapons bay doors collapse.

The extended bays also enabled integration of the ALOSNW stores for the FSAV

requirement, resolving design driver 3. The impact on OML was a high wing position

as seen on both the F-35 and F/A-22A which had implications for the undercarriage

and wing internal structure.

Proposed resolution of Design driver 4:- The size of the crew compartment for

initial sizing was evaluated to accommodate a volume of 7.5m3 to permit two ejection

ACES II seated crew, one HUD, and one console this resulted in length of 4.35m

width of 1m and a height of 1.2m, from the volumes given in table 2 below for GFE,

this resulted in an increase over the initial size of the F-35 single crew compartment

shown in figure 20, which was estimated at 0.7m width, 1.1m high, and 1.45m long

for this study. The ejection seat pitch and slope angle shown in figure 21

accommodated both crew members with full TLSS tactical life support system suits

(Nuclear / Biological / Chemical and g suits) and was identical to that of the F/A-22A,

and proposed for the F-22B and was designed to give a 99 percentile crew a clear

ejection trajectory, good situational awareness (-150 over the nose, and -45

0 over side

pilot vision angles), access to all interactive control systems MFD‟s, side-stick flight

controller, etc, and good ride quality. This extension in the forward fuselage resolved

design driver 4 and was common to all five configurations evaluated. Like the F/A-

22A noted above in design driver 2 the canopy transparency, of the FB-24 was to be a

3/4 inch thick fusion bonded and drape forged Sierracin Sylmar Corporation unit of

tinted monolithic polycarbonate, 194 inches in length, 45 inches wide, and 27 inches

high, with a rear sloping canopy bow for support in advent of bird strike, this is similar

to that proposed for the F-22B trainer shown in figure 22. The canopy shares the rear

hinge opening method of the F/A-22A, and not the forward hinge currently selected

for the F-35 family, this is due to the weight and size of the canopy, which in

emergency ejection would be separated to clear the aircraft with the aid of the airflow

over the forward fuselage and not have to act against it as is the case with the F-35.



32

Figure 20:- F-35A first CTOL SDD aircraft front fuselage nearing completion at

Lockheed Martin Georgia facility. Source Lockheed Martin F-35 general release.

Figure 21:- The ACES (Advanced Concept Ejection Seat) II which has been

proven in the F/A-22A, and therefore would only require minimal compatibility

testing for use in the FB-24. Installation was at 770 tilt back angle to the cockpit

floor. Source: - Reference 14.

770



33

Figure 22:- The proposed but cancelled training variant of the F/A-22A namely

the F-22B, which could form the base for the Advanced Regional Bomber, and

serves as the basis of the FB-24 canopy design. Source: - Reference 14:- F-22

Raptor (America’s Next Lethal War Machine) by Pace. Steve: part of the Walter J.

Boyne Military Aircraft Series Volume 1: Published by McGraw-Hill 1999.

Proposed resolution of Design driver 5:- This study proposes the use of the F-35C

undercarriage, actuators both: - EHA actuators for the primary flight controls and

rotary actuators for the leading edge flaps, avionics and mission systems (with the

exception of the current radar) to reduce risk and maintain commonality, because the

SDD F-35 programme will validate these systems ahead of the AIA / FB-24 concept

demonstrator PDR (2010), enabling the systems experience to be incorporated at this

early stage. This study also proposes the use of the ACES II proven F/A-22A ejector

seats detailed above.

The F-35C fuel system with a hose and drogue flight refuelling system, instead of the

current USAF system of flying boom refuelling (which is currently under review)

reducing the fuel system complexity. The proposed engines are based on flight proven

and near flight designs namely the F-136 and the YF-120 both offer lower risk then a

totally new engine design and development cycle.

The current design would use the same signature reduction technologies employed by

the F/A-22A, and F-35 aircraft explored in appendices B, but would be readily

adaptable to concepts after these aircraft enter service and mature. Also innovations in

manufacturing which were not at a high state of maturity for these aircraft could be

employed on the FB-24 and these will be covered in section four of this thesis.



34

Table 2:- Government Furnished Equipment.

Item Volume, ft3 Weight, lb Cost, K$

Avionics

• Base Suite

- ICNIA1 3.0 100 200

- 3 x MFD‟s 1.5 20 60

- Head-Up Display 1.6 35 20

- Data bus 0.5 10 10

• ECM Equipment

- INEWS2 3.0 100 500

Flight and Propulsion Control System

Vehicle Management System 1.0 50 200

Fire Control Systems

• IRSTS3 2.0 50 300

• Active Array Radar 6.0 450 1000

• LANTIRN Navigation System 3.0 350 200

• LANTIRN Targeting System 2.5 350 100

• HARM Targeting System 3.0 150 100

Systems and Equipment

• Electrical System (one engine) 1.0 220 40

• Auxiliary Power Unit (APU) 2.0 100 50

• Ejection Seat 8.0 160 100

• OBOGS5 1.0 35 10

• OBIGGS6 1.0 35 10

Air-to-Air Weapons ASRAAM Missile Launch weight: 192 lb Length: 9.51 ft Max span: 1.6 ft Body diameter: 0.54 ft Launcher rail weight: 50 lb Launcher rail length: 9.5 ft M61A1 20 mm Cannon Cannon weight: 275 lb Length: 74 in Max diameter: 10 in Ammunition feed system (500 rounds) weight: 300 lb Ammunition drum length: 25 in Diameter: 25 in Ammunition (20 mm) 0.58 each Returned casings 0.26 each

Air-to-Ground Weapons JDAM GBU-31PIP Release weight: 2,115lb Length: 12.38ft Max span 2.1ft Cost $20,000 ALOSNW (Projected weapon for FSAV)

7

Release weight: 2,500lb Length: 14ft Max span: 1.5ft Warhead: Evolved W69 (200kts)



35

1

Integrated Communication, Navigation, and Identification Avionics 2

Integrated Electronic Warfare Systems 3

Infrared Search and Track System with laser ranging 4

Synthetic Aperture Radar 5

Onboard Oxygen Generation System 6

On-Board Inert Gas Gen. System 7 The ALOSNW is system developed by the author and not a current weapon.

(Reference 1(b): - Aero Engr – 482: Dr Brandt: USAF Academy 1999: reference 15:-

RAF Equipment fact sheet 13 ASRAAM)

For this study the F-110-132 engine is used for integration in the structural layout

because the physical dimensions and weight figures are in the public domain and were

comparable to both the F-136 and YF-120 VCE engines proposed for the FB – 24.

Figure 23 provides a reference of current fighter and strike aircraft loadings which are

expressed as follows:- Thrust loadings are expressed as a ratio of static thrust against

weight and are an indicator of specific excess power during hard manoeuvring, and

also acceleration and initial rate of climb (thrust loadings greater than 1:1 are common

in modern fighters): Wing loading, given in lb/ft2 (kg/m

2) are an indicator of

instantaneous turning ability; combined with thrust loading these indicate sustained

turning ability.

Figure 23: - Comparison of wing and thrust loadings for modern strike aircraft.

STRIK AIRCRAFT WING AND THRUST LOADINGS.

0

0.2

0.4

0.6

0.8

1

1.2

1.4

0 50 100 150 200 250

W/S (lb/ft^2)

T/W

(lb

/lb

)

Series1

Tornado IDS

F-35B

F-35A

F-35C

Jaguar GR.1AF-117A

EF 2000

Tornado F2

F-16C

F/A-18EF-14D

Mirage 2000-5

JAS 39A

Rafale

F-15E

F/A-22



36

3.2.1 Configuration design challenges:

Before detailing the conceptual designs studied for the FB-24 manned and A-24

UCAV derivative it is instructive to highlight the design challenges faced in designing

a supersonic and supercruise capable aircraft, which are very different from those of a

subsonic aircraft, as this will show the reasoning behind the design decisions taken in

the subsequent sections of this thesis. The supersonic military aircraft is essentially

two aircraft in one, optimised for the supersonic mission but also requiring

satisfactory subsonic characteristics to allow safe and reasonable handling at low

speeds. Because the physics of supersonic flow are very different from the physics of

subsonic flow every good supersonic aircraft design is of necessity a careful blend

and compromise between good supersonic and good subsonic characteristics.

The major the major design challenges pertaining to the design of the FB-24 and A-24

were as follows:-

1. Shock waves, which occur at transonic and supersonic speeds, cause a large

increase in drag – supersonic wave drag. Supersonic aircraft need to be

designed with slender bodies (high finesse ratio) and thin wings with sharp

leading edges to reduce wave drag. These same features are not good for

subsonic speeds.

2. Careful consideration must be given as to where the shock waves occur on the

aircraft, and where they impinge on the aircrafts surface. Shock impingements

can cause flow separation and local hot spots of intense aerodynamic heating,

with resulting structural implications.

3. The centre of pressure (therefore the aerodynamic centre ac) for the aircraft

shifts dramatically aft when the aircraft accelerates from subsonic to

supersonic flight. This creates a major design challenge for maintaining the

stability and control characteristics of the aircraft.

4. Airframe – propulsion integration was a design challenge which the author

feels has been resolved by adoption of the current F-35 family intake

configuration which has been flown at Mach 2, but this requires further

aerodynamic analysis which is beyond the scope of this thesis.

5. Aerodynamic heating, which is usually negligible for subsonic aircraft design

became an important design consideration for the FB-24 and A-24, as

aerodynamic heating increases approximately as the cube of the velocity. This

not only has material selection and structural implications, but also has

implications on the infra - red signature of the aircraft, especially in the case of

the wing leading edges and intake lips, and with a desired supercruise speed of

Mach 1.6 this more of a challenge for the FB-24 than on the F/A-22A.

These were the major aerodynamic design challenges faced in the design of the

common airframe for the FB-24 and the A-24 aircraft, which are common to most

other supersonic military aircraft depending on the duration of the time spent in

supersonic flight regime. Attempts to meet these challenges are detailed below in the

following sections.



37

3.2.2 FB-24 and A-24 Fuselage configuration design selection:

The common fuselage design was intended to achieve the best streamlined shape with

the minimum surface area and highest finesse ratio capable of supercruise

performance, and lowest weight which was, capable of satisfying the internal volume,

and commonality requirements with the F-35C of the FB-24. Both the drag and

weight of the fuselage are significantly influenced by the surface area, and the grater

the surface area the higher the values of drag and weight become. General

aerodynamic considerations are given in references 4 and 7 which have been adopted

where there was no conflict with the low signature requirements of the aircraft. The

following general approach was adopted: - steps and gaps were avoided: gradual

changes in cross-section: shaping and treatment of all unavoidable protuberances e.g.

blade antennas, air data probes: fairing in the canopy as described above: the

positioning of the maximum cross-section to minimise wave drag due to

compressibility: use of area rule to eliminate the formation of compression waves

which would otherwise originate at the wing root to turn the flow parallel to a straight

fuselage, as Mach number increases (detailed below): and the smooth reduction of the

tail configuration from maximum cross-section to nominally zero, with consideration

being given for the clearance angle between the main undercarriage and the tail of

120.

For supersonic and supercruise capable aircraft, a major consideration in the design is

the distribution of the total volume along the length of the aircraft of with the fuselage

normally being the major contributor to this volume. The distribution was analysed

for both the F-35C and YF-22 using published data, and analysis models

commensurate with conceptual design as detailed below, to enable a target

distribution for the FB-24 to be determined. Ideally the total volume distribution

along the length should be smooth and, somewhat simplistically, close to a sinusoidal

shape in order to minimise volume wave drag, similar to that for the YF-22 shown

below in figure 24.

Figure 24: - YF-22 Cross-sectional area plot using the Jet 306, showing the

overall sinusoidal shape of the volume distribution over the aircraft length,

measurements from 1/72 scale drawings:- Reference 16:- The ATF Contenders:

YF-22 & YF-23 Air Superiority into the 21st Century 3

rd edition: by Sun Andy:

Published by: - Concord Publications Ltd Hong Kong 1991.



38

The overall length to effective diameter ratio or finesse discussed in section 3.1.2, is

also of fundamental importance, the higher it is the lower the wave drag, for this

initial work measurements were made from public data of the maximum cross-

sectional area (reference 4) including the intake area from which an equivalent

diameter was calculated for the YF-22 and the F-35C because neither aircraft has a

circular cross-section, and the published respective fuselage lengths were divided by

their respective equivalent diameters as follows: -

YF-22 overall fuselage length = 19.60m (64ft 2in) / core maximum fuselage

diameter = 2.64m (8ft 8in) gave a finesse ratio = 7.42:

F-35C overall fuselage length = 15.69m (51ft 5in) / core maximum fuselage

diameter = 2.41m (7ft 11in) gave a finesse ratio = 6.59.

Therefore for the FB-24 to be supercruise capable a finesse ratio (in combination with

major drag reduction measures covered above) closer to 7.42, of the YF-22 instead of

the 6.59 for the F-35C was required and this could have been approached by retaining

the core fuselage maximum diameter of 2.41m (7ft.11in) and increasing the fuselage

length to 18.89m (61ft.11 1/2in), this gave the finesse ratio of 7.83 from: -

FB-24 overall fuselage length = 18.89m (61ft 11 1/2in) / core maximum fuselage

diameter = 2.41m (7ft 11in) gave a finesse ratio = 7.83.

Although basic assessment indicated that this overall length would could be reduced

to 18.00m and would still obtain a finesse ratio of 7.47 which was closer to that

calculated from the available data for the YF-22 from: -

FB-24 overall fuselage length = 18.00m (59ft) / core maximum fuselage diameter

= 2.41m (7ft 11in) gave a finesse ratio = 7.47.

The fuselage length could be increased further to 18.29m (60ft) without a high

corresponding increases in weight but would require a modest increases in empennage

control surface area over that required for a length increase to 18.00 (59ft) to control

pitch stability. Maintaining the original maximum cross-sectional area while

increasing the fuselage length to 3.05m (10ft) enabled retention of the original F-35C

weapons bay width based on the CATIA V5 F-35C 230-5 OML model, the result of

increasing the fuselage length to 18.29m (60ft) was calculated to give a finesse ratio

of 7.58 from: -

FB-24 overall fuselage length = 18.29m (60ft) / core maximum fuselage diameter

= 2.41 (7ft 11in) gave a finesse ratio = 7.58.

This 3.05m (10ft) extension was considered the optimum fuselage growth the FB-24 /

A-24 study, theoretically for a fixed internal volume subsonic drag is minimised by a

finesse ratio of 3.0, while supersonic drag is minimised by a finesse ratio of about 14

with most aircraft falling between these values. Ideally the local cross-sectional area

of the fuselage should be matched to that of the other volume contributions from the

wing and empennage, to give a smooth overall volume distribution.



39

Although this could not have been achieved precisely until the overall conceptual

layout of the aircraft OML had been determined consideration was given to this

requirement from the start of the fuselage design.

The wing – fuselage junction is another critical area requiring careful design to reduce

drag, and decisions taken at this initial layout stage had a major impact on the rest of

the design. The shoulder wing position shown in figure 26 was adopted with the wing

above 65% of the fuselage side wall, this is similar to the F-35C and slightly lower on

the fuselage than the F/A-22A, but unlike the F-35 family the wing was not a single

unit and was joined at the spar ends to the fuselage frames as with the F/A-22A, F-16,

etc, this enable the frames to be single piece and simplified the integration of the

weapons bays. The shoulder wing position was considered to have the least drag, and

introduce the least problems for undercarriage design, compared with wings which

pass totally over or under the fuselage shown in figure 25, e.g. the Boeing X-32,

Harrier, (both restricted by powerplant installation), SEPECAT Jaguar, or the BAE

Systems Hawk and T-45, all requiring large fairings to ensure satisfactory airflow.

(a) High – wing F/A-22 and F-35 JSF.

(b) Mid – wing F-16 / Rafale / Gripen.

(c) Low – wing Eurofighter Typhoon.

Figure 25:- Sketches for the comparison of (a) high – wing, (b) mid – wing and

(c) low – wing configurations with frontal views from Reference 17:-

Superfighters (The next generation of combat aircraft): by Williams. Mel:

Published by: - Airtime Publishing: 2002.

Anhedral

Dihedral



40

(a) High – wing position: - This position for the wing is commonly found on large

civil and military cargo transports and small commuter airliners such as the BAe 146,

and enables the fuselage to be placed lower to the ground which simplifies loading in

the case of transports. This configuration enables the whole fuselage section to be

used for cargo stowage without interruption from the wing box, passing through the

cabin this was partly why the Lockheed C-5 was selected over the Boeing 747 as the

USAF‟s heavy lift cargo aircraft.

The high – wing also has greater lateral, rolling stability. For low wing configurations

as in figure 25(c) a dihedral upwards slope is usually built into the wing to increase

lateral, rolling stability.

The reason for this is that when an aircraft rolls the lift vector tilts away from the

vertical, and the aircraft starts to sideslip in the direction of the lowered wing. Where

a dihedral has been incorporated into the wing design of the extra flow velocity

component generated by the sideslip creates an increasing lift on the lowered wing,

hence tending to restore the wings to a level equilibrium position, and this is the basis

of lateral stability for naturally stable aircraft design of low wing aircraft. High – wing

aircraft on the other hand are much more stable in this regard requiring no dihedral,

this is because the extra flow velocity component generated by the sideslip when the

aircraft rolls creates a region of higher pressure in the flow interaction region between

the fuselage side and the bottom surface of the lowered wing at the wing root. This

increased pressure under the lowered wing has the effect of rolling the wings back to

the level equilibrium position.

Indeed the high wing position can be too stable in roll and has to be countered to

improve aircraft manoeuvring as in the case of the Lockheed C-5, C141, and BAe

146, and Harrier where an anhedral or downward slope shown in figure 25(a) was

adopted to reduce over stability in roll manoeuvres.

(b) Mid – wing position: - The mid – wing position used on the Lockheed U-2 spy

plane and the F-16 jet fighter shown in figure 25(b), was originally favoured by the

author because of its low drag, due to the fact that of all three options the mid – wing

configuration has the minimum wing – body interface, and unlike the high – wing and

low – wing positions requires no fillet to decrease wing – body interference, and

neither anhedral or dihedral for stability refinement. However the mid – wing has a

major structural disadvantage namely the bending moment due to wing lift must be

carried through the fuselage, and unlike the case of high – wing, and low – wing

positions where the wing torsion box can be extended across the fuselage, the mid –

wing requires heavy ring frames attached to the leading edge, and trailing edge, and

intermediate wing torsion box spars. The more spars running the length of wing

results in a greater number of fuselage ring frames carrying the bending moment

across the fuselage in the in the case of the F-16 four wing panel root attachment fish

plates form the spar to frame interface on each side of the fuselage, where four frames

are attached to nine spars in each wing. In the case of the high aspect ratio (10.6) U-2

which had only three spars, the bending moment was distributed trough twelve wing

attachment joints (six each side of the fuselage) which mated the wings to the wing

root attachment ribs and hence to four main fuselage frames which carried the

bending moment across the fuselage.



41

These heavy carry through frames add considerably to the empty weight of the

aircraft and mid - wing position was considered to be too heavy and complex for the

FB-24 and A-24 configurations.

(c) Low – wing Position: - The major advantages of the low – wing position over

both high – wing, and mid – wing positions is in reducing undercarriage physical size

and weight, and when retractable undercarriage is considered this can be retracted into

the wing torsion box, which for most aircraft is the strongest component of the

airframe structure.

Although for both the high - wing and mid – wing can employ a centreline bicycle

undercarriage for retractable units. Additionally the low – wing configuration requires

dihedral for lateral stability as shown in figure 25(c) and a fillet at the wing body

interface to minimise drag inducing aerodynamic interference, and the reasons for

filleting are covered below.

Until the early 1930‟s most mono-planes were of the high – wing configuration which

was largely the result of the aerodynamic interference at the wing to fuselage junction

which was found to be worst in low – wing configuration. Putting a circular fuselage

on top of the wing has the effect of producing a pair of rapidly diverging surfaces

which steepens an already adverse pressure gradient almost guaranteeing flow

separation and inducing drag, and reduced lift, furthermore the separated flow could

impinge on the empennage horizontal tail resulting in further stability and control

problems for the low – wing configuration (covered in detail in reference 13). It was

only through the discovery of the beneficial aerodynamic effects of mounting a fillet

at the wing body junction at the California Institute of Technology CalTech in the

USA that largely overcame these flow separation problems that enable the low – wing

configuration became widely adopted in modern aircraft designs, and what reaming

inferiority there was in the low – wing configuration could be addressed by dihedral

slopping for roll stability and was compensated for by the reduction in undercarriage

length and weight and the ground cushioning effect on landing with the Wing In

Ground WIG effect.

Figure 26:- Shoulder mounted wing position of the F-35 family adopted for the

FB-24 and A-24 airframes leaving clear weapons bay each side of the fuselage,

ferry tanks would be carried by both the FB-24 and A-24, as is the case with the

F/A-22A.

Unobstructed corner weapon bays.

Shoulder mounted wing. Shoulder mounted wing.

Under - wing clearance for

stores pylons.



42

The wing setting on the fuselage at this stage of the configuration was such that the

fuselage was horizontal and the wing had a 0.250 positive angle of attack in common

with the F-35 C, this was a relatively low angle of attack at the bottom of the common

range of 00 – 4

0. The wing root to fuselage joint was a smooth continuous curvature

blend accommodating both aerodynamic and stealth requirements.

For the initial configuration the common fuselage the length was chosen as 18.59m

(61ft) and the core diameter without intakes was chosen as 4.161m (13ft 8in). The

term common fuselage in this thesis denotes that for all wing and empennage

configurations the forward fuselage was identical, and for the trapezoidal, and

cranked arrow wing planforms either the four tail or twin tail empennage aft fuselage

configuration was used, which were created as interchangeable units, for the delta

only the twin tail aft fuselage unit was used. The centre was modified to interface with

each of the three wing planforms studied.

As stated above the common fuselage needed to reduce the peak transonic drag rise,

although currently no closed-form analytical formulas exist to predict the transonic

and even computational fluid dynamics which has been applied the computation of

transonic flows for more than 25 years, dose not always give the right answer,

principally due to uncertainties in the calculation of shock induced separated flow.

There are however two principle design features have been developed reduce the drag

rise itself or delay its effect, namely: transonic area-rule and supercritical aerofoil

(covered in the next section) it is worth using the example given in Dr Anderson‟s

work: - Aircraft Performance and Design (Reference 18: - Aircraft Performance and

Design: by Dr Anderson. D. John. Jr: University of Maryland: Published by: -

McGraw-Hill: International Addition 1999) as a demonstration of the benefits of

transonic area-ruling. Before the start of the 1950‟s designers believed that the way

forward in obtaining high Mach number performance lay in fuselage OML‟s based on

rifle bullets like that of the X-1, and Miles M-52, and they did not realise that kinks in

the cross-sectional area distribution where the wing was added to the fuselage caused

a large transonic drag rise. However by the mid 1950‟s, experimental work conducted

by Richard Whitcomb (based on his personal intuition) who was an aerodynamicist at

NACA (now NASA) Langley Aeronautical Laboratory demonstrated that contrary to

earlier opinion the addition of the wing had a profound effect on the performance of

transonic and supersonic aircraft. These results showed that the best performance

could only be obtained where the cross-sectional area distribution was smooth, this

could be achieved in part by decreasing the cross – sectional area of the fuselage in

the wing region to compensate for the cross-sectional area increase due to the wings.

The resulting fuselage OML has been likened to the old American Coke-a-Cola bottle

of the 1950‟s in shape, and pictures of pre area ruling (YF-102 prototype) and post

area ruling (F-106) on aircraft are shown in figures 27(a) and (b). Figure 27(c) shows

a schematic representation of the effect of area in reducing the peak transonic drag

rise. The first application of area-rule was to the YF-102 when flight testing

demonstrated that the prototype with a bullet shaped fuselage could not exceed Mach

1, this resulted in a major redesign effort which incorporated Whitcomb‟s area rule to

reduce transonic drag. The modified YF-102A achieved supersonic performance in its

first flight providing the USAF with its first operational supersonic delta wing fighter

subsequently this technique was adopted on all supersonic military aircraft.



43

Figure 27(a) Non area ruled F-102. Figure 27(b) Area ruled F-106

Figure 27(c) Schematic of the drag – rise properties of an area-ruled and non-

area-ruled aircraft.

CD

1.0 M 0.0

Without area rule

With area rule

Wasp Waist

Area Distribution. Amax = 41 ft2Area Distribution. Amax = 45 ft2

Note: Both aircraft have the same internal volume



44

Another feature of the fuselage which had both aerodynamic and signature reduction

benefits were the chines which ran from the radome to the end of the aft fuselage, and

which decreased the travel of the neutral point as M is increased which was

important for static longitudinal stability where the neutral point must be located

behind the aircrafts centre of gravity. The normalised distance between the centre of

gravity (cg) and the neutral point is called the static margin, this is positive when the

neutral point is behind the centre of gravity, and static longitudinal is achieved, and

hence the higher the positive static margin value the more stable the aircraft will be.

However, too much of a positive static margin is not desirable, because the aircraft

will be too stable for manoeuvring requiring a large elevator deflection to trim the

aircraft because of the large distance between the neutral point and the centre of

gravity. The net result is a trim drag which is unacceptably large. When cg is

positioned to give a good subsonic static margin it will be too large at supersonic

speeds, and conversely if the cg is located to give the correct static margin at

supersonic speed, the subsonic static margin is likely to be negative (with the neutral

point ahead of the centre of gravity), thus making the aircraft unstable in subsonic

flight. The chines offer a potential solution and the chine concept has been used on the

SR-71, YF-23, and a lesser extent on the F/A-22, and F-35 so empirical application

support this view. The additional aerodynamic benefit observed on both the SR-71,

and YF-23 at supersonic speeds was the increased directional stability (yaw stability)

imparted by the chines. For example consider fuselage of cylindrical cross – section at

a small yaw angle to the flow, this body will experience cross flow separation with a

resultant large side force, as shown in figure 28(a). However when the fuselage is of a

blended chine cross – section at the same small yaw angle to the flow no separation

occurs and the flow remains attached subjecting the body to a much reduced side

force, shown in figure 28(b). Therefore the chines are beneficial imparting additional

directional stability with the consequence that the vertical tails can be reduced in size

or combined with the elevators to produce ruddervators, resulting in reductions in

weight, skin – friction drag and systems complexity.

Figure 28: - Schematic showing cross flow streamlines over: - (a) a cylindrical

body: and (b) a blended body with chines.

Round

body

Blended

chine-body

Flow

separation.



45

The same common fuselage was slated for the A-24 UCAV variant of the FB-24

which was analysed with the FB-24 in this thesis, as an external configuration and

initial structural layout conceptual study only.

3.2.3 FB-24 and A-24 Wing and Empennage configuration design selections:

The wing is the most important major component of an aircraft as it gives the aircraft

its characteristic external appearance and determines its flight characteristics. The

many variously configured wing shapes available emphasises that to all intents and

purposes the ideal wing dose not exist and that like all aircraft design wing design is a

compromise where a given wing design will possess specific properties which can

only be attained with that configuration, while the other performance requirements are

met to a grater or lesser degree, the preference for a particular wing characteristic

having been laid down in the design requirements. Wing design has been progressing

as with all other areas of aircraft technology and this development is on going for both

military and civilian aircraft. To aid the reader in the rest of this section the author

will give a brief overview of these developments based on Reference 19:- Modern

Combat Aircraft Design: by Huenecke Klaus: Published by Airlife Publishing: 1987

as follows:-

During the early 1940‟s the general wing shapes for piston engine aircraft of the

period were dominated by the elliptical and trapezoidal wing planforms of relatively

high aspect ratios. The Spitfire typifies the elliptical wing with curved leading and

trailing edges, while the Messerschmitt Me 109 and P-51 Mustang were typical

examples of the trapezoidal planform with straight leading and trailing edges. These

wing shapes were optimal for the speeds which could be attained by these aircraft

which was in the order of 700km/h (435mph) corresponding to Mach = 0.55, because

of their favourable drag characteristics, as shown in figure 29 below. With the

introduction of the turbojet engines in the late 1940‟s, and the greater speeds these

made possible the limitations of these early wing planforms became apparent, and in

order to achieve the speed potential of the new engines the transition to swept back

wings became essential for high subsonic aerodynamics. The first application of wing

sweep was in the Messerschmitt Me 262 with a slight sweep back of the leading edge

by an angle of 18.50 and the radical rocket powered Messerschmitt Me 163 with a

wing sweep of about 500, swept wings were not a feature of the first allied jets the

Meteor and the P-80 Shooting Star which had classical straight wings.

By the 1950‟s the USA and the USSR had absorbed the original German research and

the North American F-86 Sabre and the MiG-15 went on to confirm the principle of

the swept back wing (shown in figure 29). In fact the F-86 it became possible for the

first time to just exceed the speed of sound in a shallow dive. Over this decade as

engines became more powerful operational speeds dramatically increased and by

1960 combat aircraft routinely flew at twice the speed of sound Mach = 2.0, and as

operational speeds increased other planforms were found to be more suitable, for

example the delta wing on the MiG-21 and Mirage, the cropped delta of the F-4 and

the trapezoidal wing of the F-104 Starfighter.



46

Figure 29: - Evolution of the modern combat aircraft wing planform source

reference 19.

By the 1970‟s the new high speed wings were found to be difficult to fly at low

speeds for landings, and landing speeds were becoming unacceptably high for the

Mirage 111 delta for example the landing speed was 200mph, and in Europe the

venerability of long airfields needed to cope with these factors was becoming

apparent also the aircraft carrier operation of new larger fighters was becoming

difficult due to high landing speeds. In an effort to overcome the low speed handling

problems variable sweep wings were developed which could translate from a high

speed low aspect ratio delta planform to a low speed high aspect ratio planform as

required imparting greatly improved low speed performance, however these have a

high weight and complexity penalty, examples included the General Dynamics F-111,

the BAE Systems Tornado, Rockwell B-1 bomber, and the Su-24. Also in this decade

aircraft were developed to fly for extended periods at three times the speed of sound

Mach = 3 i.e. the Mig-25 Foxbat and SR-71 Blackbird see figure 29 above.



47

While the MiG-25 featured a conservative swept back wing which had a clearly

separated wing and fuselage configuration, the SR-71 exhibited a delta-like wing

which was blended into the fuselage both reducing drag and RCS.

During the 1980‟s the demand for grater manoeuvrability for close combat led to the

development of hybrid planform wings composed of several simple planforms, such

as in the case of the F-16 and F-18 above, where the main wing is trapezoidal with the

leading edge swept back at between 300 and 40

0 and with a very slim delta portion

installed in front as a strake or leading edge extension (LEX).

In the early 1990‟s the development of thrust vectored nozzles, and advanced fly by

wire control systems enabled high agility from a verity of unstable and stealthy

planforms the most extreme example being the F-117 Nighthawk who‟s multi facetted

design is completely unstable, however the latest 5th

generation United States fighters

the F/A-22A and F-35 appear quite pedestrian compared with the European Typhoon

and Rafale but the former rely on advanced stealth shaping, materials, and thrust

vectoring to make the dog-fights the European aircraft are designed for largely

unnecessary, figure 30 shows a comparison of YF-22 with Eurofighter Typhoon.

Figure 30:- YF-22 and Eurofighter configuration comparison.

YF-22

Eurofighter



48

As outlined above in sections 1.1 and 3.2, three planforms with high lift devices were

evaluated in this thesis. In order to obtain the best initial configurations to go forward

into the Jet306 study phase the wing designs were formulated on the basis of three

study areas:-

3.2.3.1:- Aerofoil section, including the use of high lift devices:

3.2.3.2:- Planform shape and geometry:

3.2.3.3:- Empennage shape and geometry.

Basically: - the first study area was undertaken to obtain the best compromise

between all of the aerodynamic, structural and mission requirements for the aircraft:

the second study area was undertaken to determine the optimum shape governed

primarily by the requirements of high Mach flight conditions (need to supercruise),

which was also influenced by aerofoil shape: and the third study area was conducted

to the optimum overall size based on operational requirements for given values of the

first two items.

Before starting the detailed study of these three areas it is necessary to discuss the

characteristic parameters for aerofoil and wing aerodynamics, starting with the

characterisation of aerodynamic forces and moments. Fundamental aerodynamics

show the motion of air around the aircraft produces pressure and velocity variations

through the flow, and although the viscosity of the medium is a fluid property, which

acts throughout the flow field, the viscous forces acting on the vehicle depend on the

velocity near the surface as well as the viscosity itself.

The pressure forces which act normal to the surface and the shear forces which act

tangentially to the wing surface are the result of the motion of the air around the

aircraft and are illustrated on a NACA 0006 aerofoil section in figure 31, below.

These pressure and shear forces can be integrated over the surface, on which they act

in order to yield the resultant aerodynamic force (R), which acts at the centre of

pressure (cp) of the aircraft.

For convenience, the total force vector is usually resolved into components. Body –

orientated force components are used for applications concerned with the aircrafts

response (for example: - aerodynamics: and structural dynamics). Consider the forces

and moments in the plane of symmetry (i.e. the pitch plane), as shown in figure 32

below, the body – orientated components are the axial force, which is the force

parallel to the aircraft axis (A), and the normal force, which is the force perpendicular

to the aircraft axis (N).



49

Figure 31: - Normal (or pressure) and tangential (or shear) forces on an aerofoil

surface.

Figure 32: - Nomenclature for the aerodynamic forces in the pitch plane.

Shear force

Pressure Force

Shear force

Pressure Force

Shear force

Pressure

Force

A

N

R

D Thrust

CP

L

U

Weight

cg



50

In the case of applications such as trajectory analysis, the resultant force is divided

into components taken relative to the flight path or velocity vector, therefore for these

applications the resultant force is divided into a component parallel to the flight path

which is drag (D), and a component perpendicular to the flight path which is the lift

(L), as shown in figure 33. For a powered aircraft its motion through the air is

determined by its weight, the thrust produced by the engine, and the aerodynamic

forces acting on the aircrafts surface. Taking the case of steady, unaccelerated level

flight in a horizontal plane (or steady level cruise), as the simplest condition the

requirements for this condition are, that the sum of the forces along the flight path are

zero and that the sum of the forces perpendicular to the flight path are also zero.

Considering only the cases where the angles are small (for example the component of

the thrust parallel to the free – stream velocity vector is only slightly less than the

thrust itself). Summing the forces along the flight path (parallel to the free-stream

velocity), the equilibrium condition requires that the thrust must equal the drag acting

on the aircraft. Summation of the forces perpendicular to the flight path leads to the

conclusion that the aircraft weight is balanced by the lift.

Consider the case common to most modern combat where the lift generated by the

wing and body configuration acts ahead of the aircraft centre of gravity as shown in

figure 33. In this configuration the lift produces a nose – up (positive) pitching

moment about the centre of gravity (cg). Therefore to trim the aircraft and make the

sum of the moments about the cg equal to zero (i.e. Mcg = 0) a force from a control

surface located aft of the cg (i.e. the empennage) is required to produce a nose – down

(negative) pitching moment about the cg, which should balance out the positive

pitching moment produced by the wing and body lift. The empennage lift force is

illustrated in figure 32 below.

Figure 33: - Moment balance required for trimming the aircraft.

cg

Weight

(W)

Lift

(L) Lift

(LT)



51

The orientation of the tail surface which produces the lift force illustrated in figure 33

also produces a drag force, which is known as trim drag, and typically ranges between

0.5% and 5% of the total cruise drag of the aircraft. Note that the trim drag is

associated with the lift generated to trim the aircraft, but dose not include the tail

profile drag (which is included in the total drag of the aircraft at zero lift conditions).

In addition to the force components acting in the pitch plane (i.e. lift acting upwards

and perpendicular to the undisturbed free-stream velocity, and drag, which acts in the

same direction as the free stream velocity) there is a side force, which is the

component of force in a direction perpendicular both to lift and drag. This side force

is positive when acting toward the aircrafts starboard or right wing.

As stated above the resulting aerodynamic forces do not usually act through the

aircrafts cg (centre of gravity). In fact the moment produced by the resultant force

acting at a distance from the origin can be divided into three components, referred to

the aircrafts reference axes (or cg) these three moment components are:- (1) the

pitching moment: (2) the rolling moment: and (3) the yawing moment, as shown in

figure 34.

Figure 34:- Reference axes of the aircraft and the corresponding aerodynamic

moments.



52

1. Pitching moment: - The moment about the lateral axis (the y axis of the

aircraft-fixed coordinate system) is the pitching moment. The pitching

moment is the result of the lift and the drag forces acting on the vehicle. A

positive pitching moment is in the nose up direction.

2. Rolling moment: - The moment about the longitudinal axis of the aircraft (the

x axis) is the rolling moment. A rolling moment is often created by differential

lift, generated by aileron and / or spoiler deployment. A positive rolling

moment causes the starboard or right hand wing tip to move downward.

3. Yawing moment: - The moment about the vertical axis of the aircraft (the z

axis) is the yawing moment. A positive yawing moment tends to rotate the

nose to starboard.

The magnitude of these forces and of these moments acting on the aircraft depends on

the combined effects of many different parameters and the major ones are listed

below:-

1. Configuration geometry:

2. Angle of attack (i.e. the aircraft attitude in the pitch plane relative to the

flight direction:

3. Aircraft size (CFD full-size model):

4. Free-stream velocity:

5. Density of the undisturbed air (hence altitude):

6. Reynolds number (as it relates to viscous effects):

7. Mach number (as it relates to compressibility effects).

The calculation of these aerodynamic forces and moments usually requires data from

a range of flow conditions rather than just the one of principle interest, therefore data

from wind tunnel test scale models as well as full scale CFD models is often used or

data from different flow conditions. In order to correlate the data for various free-

stream conditions, and configuration scales, the measurements are usually presented

in dimensionless form, and in this form, the results are independent of all but the first

two parameters listed above, i.e. configuration geometry and angle of attack. However

in practice flow phenomena, such as boundary – layer separation, shock – wave /

boundary layer interactions, and compressibility effects, limit the range of flow

conditions for which dimensionless force and moment coefficients would remain

constant. For these cases, parameters such as the Mach number and the Reynolds

number appear in the correlations for the force and moment coefficients.



53

3.2.3.1:- Airfoil selection: -

The characteristics of the airfoil section are defined by several shape parameters of

which the most significant are shown in figure 35 and include: -

1. The maximum thickness to chord ratio and its chordwise location:

2. The nose radius, which should be relatively large to give good maximum lift

coefficient CLmax:

3. The degree and distribution of camber, if employed, some degree of camber is

common for wing sections to enhance lift characteristics:

4. The trailing edge angle, which is usually made as small as possible within

handling and manufacturing constraints.

The NACA nomenclature is used to describe a wide range of airfoil sections in use

today these were developed in the 1930‟s through to the 1950‟s and below is the

descriptive nomenclature for the four digit airfoil section series:-

4 digit code used to describe airfoil shapes:

1st digit - maximum camber in percent chord:

2nd digit - location of maximum camber along chord line (from leading

edge) in tenths of chord:

3rd and 4th digits - maximum thickness in percent chord:

For example: NACA 2412 with a chord of 4 feet:

A max camber: 0.08 ft (2% x 4 ft):

Location of max camber: 1.6 ft aft of leading edge (0.4 x 4 ft):

Max thickness: 0.48 ft (12% x 4 ft).

AERO 315

Airfoil Characteristics

Mean camber line

Chord line

Chord

x=0 x=c

Max thickness

Max camber

Leading edge Trailing edge

x

z

Figure 35: - Characteristics of an airfoil section



54

Figure 36: - Airfoil forces and moments.

Figure 37: - Airfoil centre of pressure.



55

Figure 38: - Aerodynamic centre 1 illustrates how the moment changes with

location.

Figure 39: - Aerodynamic centre 2 illustrates the point on the airfoil where the

moment is independent of the angle of attack.

For the conceptual design phase the most critical design parameters for this high

speed aircraft were the maximum lift coefficient, the drag coefficient, and the

moment coefficient, which were obtained from the NACA airfoil data charts,

although consideration was given to the following characteristics as advised in

reference 4: - Aircraft Conceptual Design Synthesis: Dr Howe. Denis: Published by: -

Professional Engineering Publishing Ltd: 2002, namely: -



56

A. The maximum lift coefficient both at low and high Mach numbers.

B. The stalling characteristics where a gentle loss of lift is preferable,:

C. The airfoil drag especially in aircraft climb and cruise conditions, when the

lift to drag ratio should be as high as possible, and at high Mach numbers.

Most airfoil sections operate with the greater part of the chordwise flow in a

turbulent state but some sections are suitable for laminar flow applications,

this can be difficult to achieve for practical applications but has the potential

for drag reduction and these supercritical sections are described below:

D. The airfoil pitching moment characteristics which may be particularly

important at high speeds causing a significant drag penalty:

E. The depth and shape of the airfoil with respect to the effect on structural

design, ease of manufacture, and possible fuel storage:

F. The slope of the lift curve as a function of incidence in that it affects the

overall aircraft attitude, especially at high values of lift coefficient, such as

required on landing.

Lift coefficient: - (3.1)

Drag coefficient: -

(3.2)

Moment coefficient: -

(3.3)

Note: no dimensional coefficients!

Where: - L, D, and M are the actual lift, drag, and moment (positive nose up) acting

on the airfoil respectively, S is the airfoil reference area and c is mean chord (S

divided by the span b), V is the flight velocity, and is the local air density.

The choice of airfoil section is broadly based on the need to obtain the best

aerodynamic efficiency in the primary operating conditions of the aircraft which in

the case of the FB-24 was supercruising flight. The airfoils selected for supersonic

aircraft are often adapted from basic biconvex (symmetrical) sections to which a small

nose radius and possibly some degree of camber has been added.

2

2

2

2

2

2

VcS

MC

SV

DC

SV

LC

pM

pD

pL



57

Maximum lift coefficient (CLMAX):-

From reference 4, the maximum lift coefficient of a basic 2-D airfoil can vary over a

wide range and is heavily influenced by the nose radius of the airfoil section, and

decreasing with the nose radius. In cruising flight buffet margin considerations could

limit the usable lift coefficient to as low as 40% of its maximum value. Therefore the

desire for a supersonic small nose radius has to airfoil had to be balanced with low

speed performance for landing and loitering.

Thickness to chord ratio (t/c):-

This is an important parameter and has an effect on CLMAX and this value is heavily

influenced by structural design requirements. Under incompressible flow conditions

relatively high thickness to chord ratios of up to 0.2 are acceptable at the root of the

wing and give a good structural depth with a small profile drag penalty. The value at

the tip is typically about two-thirds of that at the root.

At higher Mach numbers in the transonic range where compressibility effects become

important, it is usual to use somewhat thinner airfoil sections and root values in the

range of 0.10 to 0.15 are common, with the tip value again being in the order of about

two-thirds of that of the root, but the spanwise variation is not necessarily linear

especially where the wing trailing edge is cranked.

For supersonic and supercruise aircraft such as the F/A-22A and the FB-24 and A-24

the need to reduce wave drag at supersonic speed is paramount therefore the adoption

of much thinner airfoil sections was necessary with thickness to chord ratios rarely

exceeding 0.06 and some as low as 0.03 to 0.02, with any spanwise variation being

very small if any.

This lead the author to select the NACA 0006 section for the wing root with a

thickness to chord ratio of 0.06000 and the NAC 64-006 section for the wing tip with

a thickness to chord ratio of 0.05813, although not as thin as some sections suggested

in reference 4, in the authors view these sections provided the right balance between

aerodynamic properties and practical structural integrity without the wing becoming

too heavy.

Critical Mach number (M CRIT) (2-D airfoil):-

When aircraft are flying in an the high subsonic flight case at or close to their critical

Mach number the rate of drag increase due to compressibility becomes unacceptable,

and a simple definition of this condition from reference 4 is that the critical Mach

number is the one at which the wave drag due to compressibility results in an

increment of 20 drag counts (0.002) to zero lift drag coefficient. It is possible

according to reference 4 to design an airfoil such that the critical Mach number is

unchanged, or even increased, as lift coefficient is increased, although this would be

achieved at the expense of a lower zero lift value of critical Mach number and it is

more usual for the critical Mach number to reduce as lift increases. Increasing the

thickness to chord ratio also results in a reduction of the critical Mach number.



58

Reference 4 presents a simple formula for evaluating 2-D airfoil critical Mach number

originally attributed to Korn (and presented by Boppe, C.W., in AGARD-FDP Special

Course notes, May 1991, equation 25), and this was used by the author for the MNCRIT

evaluation of the 2-D airfoil sections selected, an was as follows: -

M CRIT = AF – 0.1 CL – (t/c) = Af – (t/c) (3.4)

Where: - M CRIT = the critical Mach number for a given form of 2-D airfoil:

CL = the lift coefficient:

(t/c) = the thickness to chord ratio.

AF is a numerical value dependant on the design standard of the airfoil section, in the

case of old the value would be in the order of 0.8 – to – 0.85, but for modern

optimised advanced airfoils values of 0.95 are possible, and this latter value was used

by the author in this thesis. The value of Af is equal to (AF – 0.1 CL).

Hence for the purpose of this design study: -

M CRIT = 0.95 – 0.1 CL – (t/c) (3.5)

The cruise lift coefficient for a highly manoeuvrable combat aircraft in reference 4 is

0.3 therefore the critical Mach number for the 2-D airfoils considered for this

conceptual design were:-

For NACA 0006 section: - M CRIT = 0.92 – (0.06) = 0.8600 (3.6)

For NACA 64-006: section: - M CRIT = 0.92 – (0.05813) = 0.8619 (3.7)

The effect of sweep which was of major importance to this design study is covered in

the next section on wing planform.

Lift curve slope:-

The theoretical value of the lift curve slope for thin airfoils is given in reference 4 as:-

dCL / d = 2 per radian (3.8)

Practical 2-D airfoils have a rather higher value but this decreases with the reduction

in aspect ratio and sweep, which is covered in the next section, but an approximate

value for swept low aspect ratio wings could be determined from: -

dCL / d = A/[(0.32+0.16A/cos1/4){1-(MNcos1/4)2}

1/2] (3.9)

Where: - A = the aspect ratio:

MN = the fight Mach number:

1/4 = the sweep of the quarter chord line.

The lift curve slope is only marginally affected by the deployment of leading and

trailing edge flap high – lift devices, unless their deployment grossly increases wing

area.



59

Special airfoil sections: -

The general range of combat aircraft airfoils are based on the NACA - 6 series the

Lockheed Martin F-16 family used the NACA-64A204 modified airfoil section where

the second digit (4) denotes the position of minimum pressure (maximum speed) in

tenths of the profile chord, i.e. 40% of chord. The letter A (instead of the standard

hyphen) already represents a modification of the original profile so that 80% of the

upper and lower chord surface contour has been removed producing a flat upper and

lower surface. The third from last number identifies the lift coefficient at which

minimum drag was expected (this value has to be multiplied by 0.1 to obtain the

designed lift coefficient (CL = 0.1 x 2 =0.2)), and the last two digits give the profile

thickness which for the F-16 airfoil is 4%. Manufactures employ either develop their

own airfoil sections or as in the F-16 case adapt NACA profiles to the special needs of

the aircraft, and the co-ordinates of either of these airfoils are military secrets and not

in the public domain. for the purposes of this conceptual design project standard

NACA thin airfoil sections were used with a thickness to chord ratio of 6% which lay

in the general range for supersonic combat aircraft namely 3 to 7% thickness, time

and resources permitting the author would have devised his own sections, but these

were deemed adequate for the level of analysis presented here.

Alongside the classical airfoils which were originally developed for subsonic speeds,

there are further types which have been developed especially for transonic and

supersonic speeds, as shown in figure 40 below.

Figure 40:- Special airfoil sections source reference 19.

Among these was the double – wedge airfoil section used on the F-104 Starfighter,

and the parabolic airfoil section, these had the common characteristic of a sharp

leading edge which effectively fixed the shock position of the supersonic flow.

However these sharp leading edges had disadvantages in subsonic flow, separation

accrued even at low which increased drag and only slight drag-reducing suction

could form at cruising speed.



60

In the 1980‟s attempts have been to apply the newly developed supercritical airfoils to

combat aircraft wing design, because these sections provided high lift at low drag

values. However these properties were only manifest within a narrow Mach range

around M = 0.8, and at higher Mach numbers, especially in the supersonic range, a

considerably greater drag than produced by conventional sections was experienced.

This was particularly due to the large nose radius of the supercritical airfoil section as

shown in bottom section in figure 40.

This type of airfoil was considered unsuitable for the FB-24 and A-24 aircraft

configurations because of the large range of speeds which had to be covered from

sustained Mach 1.6 supercruise, BCA / BCM, to an approach speed in the order of

120kts. The principal application of this airfoil type would be large transport aircraft,

with a single design point in terms of a primary Mach number and flight altitude.

High lift devices: -

As stated in section 3.1.1 the FB-24 and A-24 employed high lift devices in the form

of a continuous leading edge plain flap, and a trailing edge plain flap, these were

incorporated into the wing design in order to obtain maximum lift at low speeds,

which could not be supplied by the clean wing.

The plain trailing – edge flap was considered a simple form of flap with an interface

which could be easily aligned and obscured using bull-nose seals on the flap and

blade seals on the wing (see appendices B), being created by rotating the rear part of

the airfoil section around one point within the section, as shown in figure 41 below.

Figure 41:- Types of trailing edge flaps, source AERO 315 USAFA Lecture notes.

The flap deflection increases the camber, so that a greater lift is created for the same

wing angle of attack. At deflections of 100 to 15

0, the flow on the upper surface of the

flap begins to separate, but this separation zone would be confined to the flap itself.

The lift is increased with increasing flap deflection and reaches its maximum value

just before the entire wing flow breaks down, when the flow separation on the flap

jumps forward from the flap on to the wing.



61

With a plain trailing – edge flap a lift increase of around CL = 1.0 would be

possible, provided that the initial section is only slightly curved or not curved at all,

and these conditions were fulfilled in the case of the selected NACA sections for the

FB-24, A-24 configurations in common with most combat aircraft. The plain trailing

edge flap represents an effective low – cost, light weight, and low risk common usage

high lift solution which was the reason for its selection for the Advanced Interdiction

Aircraft System.

Figure 42:- Lift and drag coefficient curves for wings with flaps, source AERO

315 USAFA Lecture notes.

Several devices could be used to control the boundary layer flow and generating lift

without unacceptable increases in drag. However, their effectiveness begins only at

high angles - of - attack in the maximum lift range where separation is occurs. These

leading edge devices shown in figure 43 contribute to a reduction in the negative

pressure peaks, and can permit increases in maximum lift of up to CL = 0.2.

The front part of leading - edge plain flap shown in figure 43, would be rotated

around a point within the airfoil section to increase the camber of the wing and move

the point of minimum pressure farther aft on the upper surface of the airfoil at high

angles of attack. This aft movement of the point of minimum pressure extends the

region of favourable pressure gradient and delays separation. The plain flap is simple

and robust requiring little servicing, also like the plain trailing edge it could be

integrated into the planform alignment of the wing and the interface obscured using

blade seals on the flap trailing edge, and bull nose seals on the wing leading edge.

Basic Wing Section

Wing with Flap

CL

Basic Wing Section

Wing with Flap

CL

CD

(a): Affect of trailing edge flap on CLmax. (b): Affect of trailing edge flap on CD.



62

Figure 43:- Leading - edge flaps and boundary - layer control devices, source

AERO 315 USAFA Lecture notes.

The effect of such leading – edge flaps and boundary - layer devices on the lift

coefficient is shown below in figure 44 and as can be seen the general effect is to raise

CLmax as the angle of attack increases.

Figure 44:- Effect of leading – edge flaps and boundary – layer control devices

on lift coefficient curves.

The leading – edge flaps effectiveness, ease of integration, low maintenance

requirement, and simplicity, were the reasons for its selection for the FB-24 / A-24

configuration.



63

Figure 45: - Aerodynamic characteristics of the NACA 0006 AIA wing root

airfoil.



64

Figure 46: - Aerodynamic characteristics of the NACA 64-006 AIA wing tip

airfoil



65

For a specific angle of attack and Reynolds number the following coefficients can be

determined from the NACA data charts for the two selected airfoil sections shown in

figures 45 and 46: -

Lift (3.10)

Drag (3.11)

Moment (3.12)

N.B.:- No dimensional coefficients.

Reynolds number effects:-

As the Reynolds number increases the transition from a laminar to turbulent boundary

layer flow, occurs closer to the leading edge of the airfoil, which causes more skin-

friction drag, but delays separation reducing pressure drag. At lower angles of attack

this change in the relative magnitude of skin friction and pressure drag may result in

either higher or lower total drag at higher Reynolds numbers. At higher angles of

attack, where separation and pressure drag dominate, the reduction in pressure drag

due to delayed separation generally results in less total drag at higher Reynolds

numbers. This is shown graphically in figure 47 below which shows an airfoil that for

higher Reynolds numbers has almost the same drag at low angles of attack, yet less

drag at higher angles of attack.

Figure 47:- Airfoil lift and drag coefficient curves for two different Reynolds

numbers.

10

10

CL

.01

10

CD

Cl

Re = 9,000,000 Re = 3,000,000

Re = 9,000,000 Re = 3,000,000

cSq

mc

Sq

dc

Sq

lc

m

d

l



66

The airfoil section selection for this conceptual design study was therefore concluded

with the choice of the NACA 64-006 and 0006 sections as being the best compromise

between structural depth and aerodynamic performance on the basis that the whole

wing would be within the supersonic wave cone. The use of leading edge plain flaps

and trailing edge plain flaps as high lift devices was justified in terms of their

simplicity, effectiveness ease of obscuration both in planform alignment and interface

concealment, and this was supported by their current use on the F/A-22A, F-35, and

B-2 bomber, all of which are stealthy platforms.

3.2.3.2:- Wing geometry selection:-

At supersonic speed a Mach wave is formed which surrounds the aircraft, forming a

cone with its apex at the tip of the aircrafts nose, and the angle of this cone relative to

the aircrafts longitudinal axis is known as the Mach angle (). This angle is a function

of the aircraft Mach number as follows (reference 1): -

= sin-1

(1/M) (3.13)

Where M = the Mach number of flight (1.6 from the design requirements in section 1)

Therefore the Mach cone angle () = 38.70

In order to keep all of the airframe within the Mach wave cone requires the wing

sweep back angle to be greater than 90 - 0 = 51.3

0.

Because the air flow in this region is much lower than the free stream velocity the

wave drag would be effectively reduced. The wing leading edge sweep angle selected

for there wing options were as follows:-

Option (A) Large trapezoidal wing LE sweep angle = 550:

Option (B) Arrow head wing LE sweep angle = 550:

Option (C) Swept delta wing LE sweep angle = 600.

The original option C was an F-16XL double delta but this was dropped in favour of

the swept delta based on the F-117A which enabled a larger wing to be constructed

with better RCS spike capabilities, and greater fuel volume. The key drivers for these

planforms were reduced RCS as shown in figure 48, reduced wave drag, and increase

fuel volume, with the reduction in RCS being based on values for RCS spikes

outlined by the USAFA Department of Aeronautics in Reference 20: - lecture notes

AERO-481 Lesson 12: - Survivability Propulsion Integration and Systems.

For the configurations RCS spikes were predicted as follows:-

Option (A) with a conventional four component empennage and a shaped fuselage as

shown in figure 50 and described in appendices B, 8 RCS spikes would be produced,

however with a two component ruddervator empennage this would be reduced to 6

RCS spikes.



67

Option (B) with a conventional four component empennage and a shaped fuselage as

shown in figure 50 and described in appendices B above 8 RCS spikes would be

produced this could be reduced to 4 RCS spikes with no empennage but would

present lateral stability issues, however these could be resolved with a two component

empennage although this in turn increases the number of RCS spikes to 6.

Option (C) with a conventional four component empennage or indeed a two

component empennage this wing planform creates 8 RCS with a low RCS planform

aligned fuselage, however this could be reduced to four spiked by opting for a no

empennage configuration.

Figure 48:- The relative effects of wing geometry and empennage configuration

on RCS spikes for a low observable fuselage with all apertures and doors

planform aligned. Source authors LO design presentation.

The initial sizing of these three configurations was based on the requirement for the

largest possible volume for fuel storage to be available within the airframe without

compromising structural integrity or aerodynamic performance. So using empirical

data for: - wing loading: empty mass to take off mass: fuel mass to take off mass: dry

thrust to weight: and payload ratios from published data for fighter, interdiction, and

bomber aircraft as bench mark ranges as shown in tables 5 the author was able to

produce realistic initial sizes for the wings using the target take off weight given by

the requirements in section one.

Figures 49 and 50 below show how planform alignment works and influences design

choices.



68

Figure 49:- F-35C planform alignment of RCS spikes to illustrate the behaviour

of an illuminated aircraft. Source authors LO design presentation.

Figure 50:- The contribution of fuselage and empennage alignment to RCS

reduction of an illuminated aircraft. Source authors LO design presentation.



69

Having reduced the forward direct specular returns by planform alignment of the

leading edges of the aircraft OML the surface current scattering contributions needed

to be considered at this configuration design stage. This surface current scattering

results from electromagnetic currents which build up on the skin when the aircraft is

illuminated by radar. These currents flow across the skin until they meet a

discontinuity such as a sharp trailing edge, a wing tip, a control surface, or a gap

around a removable panel or door. At a discontinuity, the currents “scatter,” or radiate

electromagnetic energy, as sure as night follows day some of this will be detected

back at the radar, this is shown in figure 51 below.

Figure 51:- Surface current scattering, source reference 20.

The surface current scattering effect is of lower intensity than the specular return, but

is still sufficient for detection, in fact a radar operator only needs two spike returns in

quick succession to obtain an attacking aircrafts direction and velocity and compute a

firing solution which at worst results in dead pilots, or at least some very unhappy

ones. The effect is strongest when the discontinuity is straight and perpendicular to

the radar beam, therefore the wing and empennage trailing edges needed to be swept

as well as the leading edges as was the case with the three planforms studied for the

FB-24 and A-24, to minimise detection from the front, note the large negative

(forward sweep) angles on the trailing edges of the wing and horizontal tails of the F-

35C in figure 49 its not just aerodynamics. Good LO design also calls for diamond or

chevron doors on every access panel as on the B-2, F-117, F/A-22A, and F-35.

The scattering of surface currents shown in figure 51 actually represents three

different types of radar returns: -



70

(1) Illuminated surface discontinuities cause diffraction, which analogous to diffracted

light through a prism, and this diffraction not only occurs at a physical edge, such as

the wing or control surface trailing edges discussed above, but at any location that has

a sharp corner this mitigated against the use of the double wedge airfoil section (see

figure 40) for the FB-24 and A-24 wing, this also occurs where there is a shadow edge

where the airfoil section curves away from the illumination source in the transition

from wing leading to trailing edge, this mitigated further against the use of a

supercritical airfoil section (see figure 40) (in addition to its high Mach aerodynamic

performance): (2) Travelling waves are another radar return product of the scattering

of surface currents and occur when a sharp discontinuity is reached and the energy

released (which cannot be destroyed) travels back to the front where it reradiates.

Both edge diffraction and travelling wave radar returns mean that for successful

signature reduction straight trailing edges perpendicular to the illuminating radar

which the original option C choice of a of a planform based on the F-16XL: (3) The

third radar return is from creeping waves which occur when the backside of the

illuminated body is smoothly curved enabling the energy to creep all the way around

the body, slightly radiating as it goes around, this cannot be directly controlled by

shaping and radar absorbent materials covered in section 4 and appendices B would

be used to suppress creeping wave effects on the FB-24 and A-24 airframes. (The

above paragraph was based on material presented by Daniel Raymer which was also

covered in reference 20: - Reference 21: - Page198: Aircraft Design A Conceptual

Approach 3rd

edition: Raymer. Daniel. P.: Published by AIAA: 1999).

Although the choice of the option (B) wing planform was founded upon a similar

planforms consideration for the Northrop Grumman / McDonnell Douglas / BAe

Mach 1.6 JAST submission, and the use of a sub-sonic version on the Northrop

Grumman / Boeing B-2 bomber, Daniel Raymer adds some further supporting

augments for this planform from both a signature and aerodynamic view – point

which are included here for completeness. As will be seen in the next sub-section a

non-tapered wing with a taper ratio of 1 shown in figure 52(a) or a diamond wing with

a taper ratio of 0 shown in figure 52(b) are aerodynamically the worst possible

configuration. The former would have excessive outboard lift especially when swept,

and the diamond wing would have insufficient outboard lift to form an elliptical lift

distribution as desired for minimum drag due to lift. However by combining these two

aerodynamically ineffective planforms as shown below in figure 52(c) and apply twist

and camber to the resulting combination, a fairly good aerodynamically efficient wing

planform would be created. The basic non-tapered wing shown in figure 52(c) is

similar to the original Northrop Grumman / Boeing B-2 bomber configuration which

was revised later on in the design cycle to the current configuration to crate a more

balanced design. The combined planform has beneficial RCS characteristics as well as

improved aerodynamic qualities in so much as when illuminated from the front aspect

the two return spikes are angles away from the receiver, and when illuminated from

the rear the trailing edge also produces two return spikes which are angled away from

the receiver as shown in figure 52, hence the four spike classification of this planform,

however when a empennage is added the number of possible reflectors is increased

rising to eight for a conventional four component tail. As can be seen from figure 48

option (B) for the FB-24 and A-24 configuration studies had a taper in the outboard

wing sections to reduce tip loading and return the lift distribution to a more desirable

elliptical form.



71

Figure 52:- Wing planform combination of planform (B):based on reference 21.

Geometric and physical dimensions of the preliminary design:

Aerodynamics Specialists were not available to contribute to the activity of planform

geometry selection and provide assistance on the basic wing geometry, or assist in any

analysis of the three planforms selected, and this was solely conducted by the author

based on reference material accumulated in the course of this project, and on the

results of the Jet306 analysis tool.

Figure 53: Wing geometry major parameters, source reference 19.

Diamond

(b)

Non-tapered but

swept wing (a)

Combined

(c)

Spike 3

Spike 1 Spike 2

Spike 4

c 1/4

S

b / 2

cr

b = full span



72

The major drivers for sizing these planforms were as follows: -

Supercruise aerodynamics:

Front and rear spar locations:

Undercarriage location to be aft of the Centre of Gravity (C of G):

RCS (discussed above):

Weight and sizing:

Usable fuel –fuel storage volume within the wing structure:

Type sizing and location of control surfaces:

Location and sizing of high lift devices.

A summary of the influence of the wing parameters given above in figure 53 on the

aircrafts performance is given below:-

1. Aspect Ratio: - The effect of increased Aspect Ratio is to improve the Lift and

Drag ratio (L/D) as shown in figure 54, and is beneficial when the useable take off

incidence is restricted by ground clearance. However, for low altitude, high – speed

fight, (as specified in the requirements document) profile drag is dominant and little

benefit is derived from high Aspect Ratio in respect of radius of action.

Lower Aspect Ratio not only reduces profile drag but gives a smoother ride to crew or

in our case Flight Systems Computer, and reduces structural loading in turbulent

conditions because „g‟ due to gust is proportional to lift slope, and the lower the slope

the more attenuated is the response.

Figure 54:- Effect of Aspect ratio on lift, source reference 19.



73

2. Sweep: - Sweep gives higher lift dependent drag and requires increased roll control

of cross wind take offs. However, it delays drag rise „M‟ and reduces lift curve slope.

In this case the sweep angle range of 55 – 60 degrees was dictated by the speed

requirements of Mach 1.6, and for options A and B the leading edge sweep angle of

55 degrees was selected and for option C 60 degrees was selected, see below for wing

sweep analysis.

3. Taper: - The taper transfers load from the wing tip towards the wing root, thus

increasing the likelihood of tip stall (which gives wing drop and pitch up on a swept

wing). Usual taper ratio values lie between 0 and 0.5. A wing with a constant profile

chord i.e. no sweep has a taper ratio of 1; a typical delta on the other hand has a taper

ratio of 0 because of its non-existent wing tip chord. By means of the taper ratio the

load distribution of the wing outwards can be influenced and a quasi – elliptical lift

distribution produced. For a swept wing increased taper gives lower trailing edge

sweep, which enhances the effectiveness of trailing edge flaps and controls (giving

reduced take off and landing speeds and improved controllability in cross winds).

4. Thickness: - Thick section wings incur a Profile Drag Penalty. Increasing

thickness, dose however, give increased maximum lift, eases mechanisation of flaps

and slats, generates a lighter internal structure and presents a greater internal volume

for fuel carriage, so there is a trade off between profile drag, mass and range.

5. Camber: - Chamber is added to enhance lift, however the configurations were

judged to have adequate lift and therefore no camber was employed on the wing, and

chamber is detrimental to low speed performance and was not really considered

worthwhile.

6. High lift devices: - These are of primary benefit on thin swept wings at supersonic

speeds. All of the wing configurations studied use high lift devices as considered in

detail in section 3.2 above.

In summary therefore sweep and low Aspect Ratio (lowest Profile Drag) give a low

lift slope, which coupled with high wing loading gives, rise to good ride quality in

turbulence.

Having considered the basic wing geometry and the wing location the aircraft C.G.

could now be discussed.

(1.) The Mean Aerodynamic Chord (MAC) may be obtained graphically from

the intersection of a diagonal line (constructed as shown above in figure 55), with

the mid chord line of the wing.

(2,) The aircraft centre of gravity can be estimated as a percentage of the MAC

for a given aircraft configuration e.g. 25 to 30% for a stable aircraft with aft tail:

40% (approx) for an unstable aircraft with aft tail as in the case of the FB-24

and A-24 configurations, and 15 to 20% for an unstable aircraft with foreplanes

such as Eurofighter.



74

Figure 55: - The wing relationship to the Aircrafts Centre of Gravity.

The effect of wing sweep on MCRIT for finite wings:-

As determined above the leading edge of the wing planforms needed to be swept in

order to keep the wing within the Mach 1.6 wave cone this also had the effect of

raising the wings critical Mach number which was additional to reducing airfoil

thickness.

The method by which this could be achieved is demonstrated in figure 56, by

considering the non-tapered swept wing, where the effective chord length increased

because the chord is measured in the stream wise direction, due to the airfoil shape the

air must flow around is a stream wise cross – section of the wing. From the geometry

shown in figure 56, the relationship between the chord of the unswept wing and the

chord of the swept wing is determined from:-

C (swept wing) = C (non-swept wing) / COS LE (3.12)

So that:-

(t max /c)(swept wing) = (COS LE) (t max /c) (unswept wing) (3.13)



75

In the cases of the airfoil sections selected for this conceptual design (e.g. NACA 64-

006) the swept wing chord expression (3.13) could be substituted into the curve fit for

M CRIT data for NACA 64-series airfoils i.e.:-

M CRIT =1.0 – 0.065[100(t max /c)] 0.6

(3.14)

There by producing an expression the critical Mach number for 3-D swept wings:-

M CRIT =1.0 – 0.065 cos0.6

LE [100(t max /c)] 0.6

(3.15)

Or for the unswept wings M CRIT:-

MCRIT = 1.0 – cos0.6

LE (1.0 - MCRIT (unswept)) (3.16)

For tapered wings as in the planforms types evaluated in this conceptual design study,

the effect was modelled by using 0.25c, the sweep angle of the line connecting the

quarter-chord points of the wings airfoils and using the maximum value of (t max /c) on the wing i.e.:-

MCRIT = 1.0 – cos0.6

0.25c (1.0 - MCRIT (unswept)) (3.17)

The effective critical Mach number MCRIT for the leading edge sweep angles used in

this conceptual design study were determined from the methodology presented in

USAFA Aero Lecture notes WIN-12 as reproduced below and Reference 22: -

Introduction to Aeronautics A Design Perspective: by Brandt S. A: Stiles R. J: Bertin

J. J: Whitford R.: Published by AIAA: 1997, in the configurations design and analysis

section 3.2.4 below.

This analysis assumes a wing with a taper ratio of 1.0, so we don‟t have to deal with

different leading edge, quarter chord, and trailing edge sweep angles. The wing as

drawing with the total Mach number over the swept wing equal to 1.0, as shown in

figure 56. Assuming that the velocity component perpendicular to the leading edge of

the swept wing speeds up as it flows over the wing by the same ratio as the flow going

over the unswept wing speeds up, so that:-

unsweptswept

unsweptswept

critcrit

critcrit

MMM

MM

M

/cos

1

cos

wingover thelar perpendicu

wingover thelar perpendicu

(3.18)



76

Figure 56:- Unswept and Swept Wings at Their Critical Mach Numbers source

USAFA Aero lecture WIM – 12 supplied within the AeroDYNAMIC V3.0

software cd Rom reference 1(b).

Then, assuming the total Mach number over the wing is 1.0, using the Pythagorean

Theorem the following equation can be derived: -

2

22

2

222

222

22

cossin

1

cossin1

/cossin11

/cossin1

unswept

swept

unswept

swept

unsweptsweptswept

unsweptsweptswept

crit

crit

crit

crit

critcritcrit

critcritcrit

M

M

MM

MMM

MMM

(3.19)

This was a relatively complex equation, but it gave useful results, and was worth the

extra complexity. Table 3 shows the critical Mach numbers this equation would

predict for a wing which has an unswept Mcrit = 0.7. Note that Mcrit does not go to 1.0

until the wing sweep goes to 90 degrees, which is a much more reasonable result.

1m

LE =450

1m



77

Table 3 Variation of Critical Mach number with Wing Sweep

MCRIT Non-Swept 0.7

MCRIT Swept Sweep Angle

0.7 0

0.70544532 10

0.721863822 20

0.749402818 30

0.787920521 40 0.836230244 50 0.890798665 60 0.944172829 70

0.984667738 80

1 90

Now, this derivation is not perfect. As Figure 56 shows, for the perpendicular

component to speed up by the same ratio as on the unswept wing, if the parallel

component does not speed up at all, then the flow must turn as it goes over the wing.

The flow does, in fact, turn as it goes over swept wings, but it turns the other way,

toward the wingtip.

However if the flow is assumed not to turn at all, then the derivation presented in

above applies:

M CRIT =1.0 - cos0.6

LE (1.0 – M CRIT (unswept))

Table 4:- Variation of Predicted and Measured Critical Mach

number with Wing Sweep based on F-111 experimentation.

Sweep Angle, deg

Mcrit, Classical

Mcrit, Modified

Mcrit, BSBW

Mcrit, F-111 Flight Test

0 0.75 0.75 0.75 10 0.76 0.75 0.75 16 0.78 0.76 0.76 20 0.8 0.77 0.76 26 0.83 0.78 0.77 0.75 30 0.87 0.79 0.77 35 0.92 0.81 0.78 0.8 40 0.98 0.83 0.79 50 1.2 0.87 0.81 0.85 60 1.5 0.91 0.84 0.9 70 2.2 0.96 0.87 72 2.4 0.96 0.88 0.95 80 4.3 0.99 0.91 90 infinite 1 1 1



78

The analytical method results were compared with those measured on actual aircraft.

For this purpose, flight test data for the F-111 with several different wing sweep

settings was used, and Table 4 compares the F-111 flight test data with values of Mcrit

predicted using the various methods described in lecture WIM-12. This gave

confidence in this prediction method over the classical method publicised in some

textbooks.

Having determined the airfoil section choice: the wing planform choices options (A),

(B) and (C): and the wing leading edge sweep angles for these planforms, the next

stage in the wing design was to size the wings in conjunction with the rest of the

airframe OML and analyse their aerodynamic properties in term of the lift and drag,

characteristics of each option and as part of the full configuration analysis was

undertaken in section 3.2.4 below.

3.2.3.3:- Empennage geometry selection: -

The principle role of the tail is to counter the moments produced by the wing. Thus

the tail size is directly proportional to the wing size and this was determined from the

moment equations examined in section 3.4. It would be expected that the tail area

divided by the wing would show a consistent relationship for different aircraft

empirical data, if the effects of the tail moment arm could be accounted for. A method

for doing this was adopted from reference 21, and was termed the “tail volume

coefficient” method for the initial estimate of tail size, which was based on the fact

that the force due to the tail lift is proportional to the tail area times the tail moment

arm.

For the vertical tail, the wing yawing moments have to be countered which are most

directly related to the wing span b W and from this relationship the “vertical tail

volume coefficient” was derived as defined in equation 3.20:-

c VT = LVT SVT / bW SW (3.20)

For the horizontal tail the pitching moments which must be countered are most

directly related to the wing mean chord C-W and from this relationship the

“horizontal tail coefficient” was derived as defined in equation 3.21:-

c HT = LHT SHT / C-W SW (3.21)

As recommended in references 19 and 21, the moment arm (L) was approximated as

the distance from the tail quarter – chord (25% of mean chord length measured back

from the leading edge of the mean chord) to the wing quarter chord. The tail moment

arm definition is shown in figure 57 below along with the definitions of tail area. Note

that the horizontal tail area was measured to the aircrafts centreline which is standard

practice according to reference 21. In the case of both the FB-24 and A-24 as part of

the low RCS design twin vertical tails were adopted and for this configuration the

vertical tail calculated represented the sum of the areas of both tails.



79

Figure 57:- The definition of tail moment arm and tail area definitions for initial

sizing.

Using historical data from Table 4.1 (Configuration data for modern combat aircraft

F-15 to MiG-25) page 59 of reference 19, an average value for the horizontal tail

volume was determined as follows: - c HT = 0.476 and a value for the vertical tail

volume was taken from Table 6.4 (Tail volume coefficient) page 125 of reference 21

c VT = 0.07 to use in equation 3.22 to determine horizontal tail areas and equation

3.23 to determine vertical tail areas respectively:-

SHT = c HT C-W SW / LHT (3.22)

SVT = c VT b W SW / LVT (3.23)

The moment arm was approximated by a percentage of the fuselage length and from

reference 21 for an aircraft with an aft mounted engine a tail moment arm range of

between 45% and 50% of the fuselage length was recommended. This was verified by

examining historical combat aircraft data from reference 19 as well as measurements

taken from public domain data for the F/A-22A and the F-35C.

For the YF-23 type ruddervator tail group configuration studies the tails were sized to

provide the same surface area as a conventional four tail group, and as stated in

section 3.2 these were given a dihedral angle of 500 based on the flight proven YF-23

aircraft.

LHT

LVT

CL

SW = wing area:

bW = wing span:

C-W = wing mean chord.

bV

bH

b/2

c1/4

c1/4

c1/4



80

Figure 58:- F-35C port – side conventional tail layout with fixed VT and all

moving HT although the flip tail was supplanted by a pivoted tail, source

authors‟ private collection.

Figure 59:- YF-23 Ruddervator all moving V – tail for FB-24 and A-24

configuration sizing the ruddervator area equalled the conventional four tail

group source authors‟ private collection.

All moving (flipping) HT

area = 5.78m2

Fixed VT (dihedral 270) total area

from rudder + tail = 4.547m2

Rudder area = 1.332m2



81

3.2.4 Common FB-24 and A-24 configuration initial sizing.

This stage of the conceptual design was intended to produce complete models of the

proposed modifications to the original F-35C aircraft to produce the definitive OML‟s

for the FB-24 and A-24 for evaluation with the Jet306 toolset, in section 3.3.

These models contain following key attributes:-

1. A common 3.048m (10ft) extended two crew position fuselage for all of the

FB-24 wing and empennage variants:

2. A common 3.048m (10ft) extended no canopy fuselage for all of the A-24

UCAV wing and empennage variants:

The following FB-24 aircraft surface models were produced:-

1. New Build 1 (NB-1):- Common two crew position fuselage with option (A)

520 leading edge sweep angle trapezoidal wing, and conventional four

component empennage:

2. New Build 2 (NB-2):- Common two crew position fuselage with option (A)

520 leading edge sweep angle trapezoidal wing, and ruddervator two

component empennage

3. New Build 3 (NB-3):- Common two crew position fuselage with option (B)

550 leading edge sweep angle cranked arrow wing, and conventional four


4. New Build 4 (NB-4):- Common two crew position fuselage with option (B)

550 leading edge sweep angle cranked arrow wing, and ruddervator two


5. New Build 5 (NB-5):- Common two crew position fuselage with option (C)

600 leading edge sweep angle swept arrow wing, and conventional four


6. New Build 6 (NB-6):- Common two crew position fuselage with option (C)



The following A-24 UCAV aircraft surface models were produced:-

1. New Build UCAV 1 (NBU-1):- Common no canopy fuselage with option (A)

520 leading edge sweep angle trapezoidal wing, and conventional four


2. New Build UCAV 2 (NBU-2):- Common no canopy fuselage with option (A)

520 leading edge sweep angle trapezoidal wing, and ruddervator two




82

3. New Build UCAV 3 (NBU-3):- Common no canopy fuselage with option (B)

550 leading edge sweep angle cranked arrow wing, and conventional four


4. New Build UCAV 4 (NBU-4):- Common no canopy fuselage with option (B)



5. New Build UCAV 5 (NBU-5):- Common no canopy fuselage with option (C)

600 leading edge sweep angle swept arrow wing, and conventional four


6. New Build UCAV 6 (NBU-6):- Common no canopy fuselage with option (C)



The intention was then to analyse these configurations against the F-35C 230-5 OML

model using the hand calculation whole aircraft analysis techniques given in reference

22 section:- 4.7 to obtain initial comparison of:- Lift: Parasite drag: Induced Drag:

and Supersonic drag, to determine which of the 6 FB-24 configurations and 6 A-24

configurations gave the highest improvement in complete aircraft drag polar over a

Mach number range of Mach 0.3 – Mach 1.6, over the baseline F-35C analysed by

the same methods over the same Mach range, and only these configurations were put

forward for Jet306 analysis, this however had to be re-scoped to meet the constraints

of this design study and a less extensive comparison was made.

To produce accurate models for initial sizing required estimates of the main aircraft

parameters and as an initial starting point the values associated with existing aircraft

of similar types to the intended Advanced Interdiction Aircraft System were

investigated. Although the A-24 UCAV was unique in function i.e. supercruise long

range unmanned combat aircraft no strictly comparable weight data exists, however

because the A-24 was basically a common airframe with the manned FB-24 the use of

manned aircraft was deemed appropriate, and if the A-24 empty weight was

significantly less than the FB-24 this would be used for additional fuel storage. To

this end a list was compiled for existing military aircraft in the: - Fighter: Interdictor:

and Bomber categories (excluding subsonic aircraft) table 5 below, from published

data using references: - 19: 23:-Jet Bombers (From the Messerschmitt Me262 to the

Stealth B-2): by: - Gunston. B. and Gilchrist. P.: Published by: - Osprey Aerospace:

1993: and 24:-Modern Fighters: by: - Spick. M.: Publishes by: - Salamander Books

Ltd: 2000.

The production of these accurate models of the study variants enabled estimates to be

made of the component weights, drag, and lift. The predictions for thrust requirement

derived below allowed down selection of the engine capable of providing the required

performance over all fight conditions. With weight, aerodynamic, and propulsion data

it was possible to perform the initial performance analysis and produce the initial

constraint diagram.



83

Table 5:- Comparison of equivalent military aircraft parameters.

Parameter Fighters radius

<700miles

Interdictors radius

1000miles

Bombers radius

>1000miles

Empty weight ratio (ME / MTO) 0.46-0.72 0.49-0.57 0.40-0.59

Fuel weight ratio (MF / MTO) 0.22-0.45 0.19-0.33 0.33-0.64

Payload ratio (MPAY / MTO) 0.21-0.28 0.09-0.32 0.08-0.134

Wing loading (MTO / S) kg/m2 261-818 225-1050 204-1194

Thrust / Weight ratio (dry) 0.65-1.48 0.44-0.56 0.26-0.52

Aspect ratio (b2 / S) 2.36-7.28 1.64-7.08 1.75-9.58

To make the data as representative as possible only three subsonic aircraft were

included in the above table studies and these were both interdictors namely the

Lockheed Martin F-117A, the British Aerospace Buccaneer S-2, and the BAe /

Boeing Harrier, all others being supersonic capable aircraft.

The highest wing loadings for all categories came form swing wing types i.e.: -

Tornado F-3 Fighter: Tornado GR-4 Interdictor: and the B1B bomber, and the lowest

wing loadings came from the Mirage 2000 Fighter: F-117 Interdictor: XB-70 Valkyrie

Bomber. The highest non variable geometry values being: - MiG-31 =666kg/m2

Fighter: Buccaneer = 550kg/m2 Interdictor: B-58 Hustler = 560kg/m

2 Bomber.

The lowest fuel weight ratios for all categories came from the: - AIDC A1 Ching Kuo

Fighter: Tornado GR-4 Interdictor: and the TSR-2 Bomber, and the highest came

from the: - F/A-22A Fighter: F-117A Interdictor: and F-111F Bomber.

The highest aspect ratios were all from swing wing aircraft understandably i.e.:-

Tornado F-3 Fighter: Tornado GR-4 Interdictor: and the B-1 bomber, the highest

values for non variable geometry aircraft were: - F/A-18E Hornet = 4.00 Fighter:

Buccaneer S-2 = 3.52 Interdictor: and A-5 Vigilante = 3.75 Bomber.

Applying this approach to the AIA system is valid for both the manned FB-24 aircraft

variant as well as the A-24 as both are conceived as extended range interdictors with

the common airframe approach for the A-24 UCAV.



84

However the AIA system dose not follow the „fighter‟ category of aircraft because of

the need to a longer distance with a heavier payload than is normal for fighters,

neither dose the AIA system fit into the „bomber‟ category of aircraft because of the

requirement for sustained high speed and lighter payload, over a modest range. The

most realistic category is long range interdictor but the current generation of

interdictors are not expected to supercruise to the targets for long periods or deliver

weapons from 45,000ft undetected, they do have greater manoeuvrability and are

expected to fight their way to and from the target at medium or low level, resulting in

a higher stressed 9g and hence heavier airframe than either the FB-24 or A-24.

From Table 5, it is clear that wide variations exist in the aircraft studied, excluding the

extreme values from variable geometry aircraft which were a niche phase in military

aircraft in the 1970‟s, it is possible to assess the variations in some design parameters,

although this only serves as a crude guide for initial sizing and the refined method of

whole aircraft drag polar analysis against the F-35C using reference 22 methods, was

used airframe refinement and for selection for Jet306 toolset submission. To enable

the selection of suitable starting values reflection on the key requirements of the AIA

system aircraft was necessary, and the following considerations raised from the

requirements: -

All but one of the aircraft in the authors survey were not supersonic in dry /

military thrust therefore the AIA system aircraft would require a higher thrust

to weight ratios than the values for interdictors and bombers.

Travelling for longer distances at supersonic speed obviously requires more

fuel than is seen in the fighter and interdictor categories in table 5 but less than

max bomber value from the Valkyrie (Mach 3).

The fuel capacity required was larger than for an equivalent sized aircraft

which favoured a larger wet wing for increased fuel storage, which unlike the

naval F-35C aircraft with its need for a folding outboard wing section, could

be wet to the tip, which could also produce bending relief.

The larger wing with the resultant lower wing loading would also help the

aircraft meet the icy runway landing requirement.

The weapons load based on the weights supplied in table 2 section 3.2 above,

and defined in section 2 constituted a relatively low payload ratio to the Max

TO weight. Where as the range to be flown at supersonic speed gave a fuel

weight ratio.

The empty weight ratio of the aircraft compared to the F-35C would be

reduced due to the de-navalisation of the airframe by removing structural

weight associated aircraft carrier landing, catapult launch loads, from

reference 21 this would be a weight saving in the order of 12%. This weight

saving could be eroded by the requirements of materials which would be

required for the aircraft to endure relatively high temperature long duration

Mach 1.6 flight, at 45,000ft like BMI composites, ceramics, and titanium.

Therefore a reduction of 10% was considered more realistic for the AIA



85

system aircraft. The use of low observable coatings (see appendices B, and

reference 13) is no longer considered a high airframe weight driver and other

methods have been adopted to eliminate structural steps and gaps (which

cannot be presented here under ITAR clearance rules) that do not

demonstratively add weight to the airframe.

Taking the above considerations into account the initial estimates for the AIA system

aircraft were made as follows and compared to calculations made for the F-35C 230-5

using OML publicised data:-

Empty weight ratio = 0.40: - this was low compared with the fighters and

interdictors surveyed in table 5 but just within the supersonic capable bomber

category and was below the navalised F-35C empty weight ratio of 0.43

with all internal and external weapons stations filled. This took into account a

non navalised 10% structural weight saving and the weight growth with

fuselage growth.

Fuel weight ratio = 0.51: - This was higher than any of the fighters or

interdictors surveyed in table 5 but was mid range for the supersonic bombers

surveyed, and was 1.8 times that of the F-35C which in its current form has

an internal fuel weight ratio of 0.27.

Payload weight ratio = 0.09: - Based on the published release weights of

ASRAAM, and JADAM – GB31 PIP weapons from table 2, and the projected

weight for the ALOSNW, and for the mission requirements weapons fit

options the actual payload weight ratio required was 0.076, however the

inclusion of the two crew and potential weapons weight growth increased this

value of 0.09. This value falls just below the lowest range points for the

interdictor range form the F-117A value and the lowest range point for

bombers from the clean F-111F.

Wing loading range = 221kg/m2 to 352kg/m: - The top of this range was

about mid range for all non variable geometry aircraft surveyed in the three

categories in table 5 and was considered the highest permissible value for all

three wing planform options, in view of the need to reduce structural weight of

the airframe. The factors influencing the final wing loading and the analysis

supporting this selection are detailed below. Compared to the wing loading

of the F-35C at MTO = 560kg/m2 the top of this range was 0.63 times the

calculated F-35C figure from model data, however the maximum vale in

the range was just above the published figure for the F/A-22A =

348.7kg/m2.(79.43lb/ft

2) the only true combat aircraft supercruise in

existence.

Thrust loading = 0.48: - This was low for fighters, in the lower range for

interdictors but higher than bombers surveyed in table 5. This was also higher

than the navalised F-35C which had a Tdry / WTO (Thrust loading) value =

0.40 calculated from published data.



86

It was now possible to use these assumed values and the fuselage length and cross-

section estimate from section 3.2.2, and the methodology of determining tail areas

from section 3.2.3.3, and the requirements data from section 2 to make the first rough

predictions of the size of the aircraft, and create accurate models for analysis.

From the requirements definition in section 2 the value of MTO = 32,213.23kg

(71,018.16lbs) a payload of 2 x ALOSNW + 2 x ASRAAM = 2442.20kg

(5,384lb) = 0.076 of MTO therefore based on a payload weight ratio of 0.09 x

MTO a further 450.98kg (994.25lb) was available to accommodate the weight

of the two crew members and weapons growth for the FB-24 manned aircraft

and for the AI system for the A-24 UCAV.

The empty weight ratio of 0.40 x MTO = 12,885.28kg (28,407.20lb) which

was below the do not exceed target empty weight set out in the requirements

detailed in section 2 of 14,141.65kg (31,177lb)

The fuel weight ratio of 0.51 x MTO = 16,428.74kg (36,219.18lb) which was

1.89 times that of the F-35C value of 8,663.61kg (19100lb).

The wing loading range of 220kg/m2 to 350kg/m

2 resulted in gross wing

areas of 146.42m2 to 92.04m

2 respectively, which were applied as target

ranges for practical wing sizing of options (A), (B), and (C). The benefits of a

low wing loading are good supersonic performance and acceleration, and rapid

roll rates, but this is at the cost of excessive speed loss during hard turns and

an increase in the speed required for maximum instantaneous rate turns as

Mach number increases as shown in figure 60 below.

The thrust loading of 0.48 x MTO (sea level static dry) = 15,462.34kgs

(34,088.63lbs) of thrust required from the engine, this dictated the selection of

the YF-120 derived F-136 (Variable Cycle Engine) engine with a

maximum thrust of 43,000lbs and a projected dry thrust of 34,400lbs over

conventional F-119 derived F-135 engine with a maximum thrust of 4300lb

but a dry thrust of only 28,810lbs, see also appendices D.

An aspect ratio between 2.00 and 3.5 was considered as reasonable for the

AIA system aircraft, and was driven by the desire to keep the wing within the

Mach wave cone, so that the shockwave drag due to volume (that part of wave

drag due to the bulk of the aircraft, which is independent of lift) was

minimised, and to keep the wing within manageable proportions with respect

to the weight of the wing, and to maintain control surface effectiveness, as the

aerodynamic centre shifts aft with Mach number.

The fuselage length and maximum cross section were determined above to

give a comparable finesse ratio to the YF-22 as length = 18.29m (60ft) and

maximum cross-section = 4.47m2.



87

Figure 60:- Effect of reducing wing loading and increasing thrust to weight

ratios on the speeds for maximum instantaneous and maximum sustained turn

rates respectively with increasing Mach number. (Chart based a diagram from

reference 22).

From the wing loading and corresponding areas calculated above and after several

iterations the most practical configurations for the three wing options were modelled

based on aerodynamics, structural and signature requirements, and these are detailed

below in figures 61 through 63 and the key dimensions were as follows:-

Option (A) figure 61:-

Total span = 13.58m: Total area = 91.76m2: Wing loading = 351kg/m

2: Aspect

ratio = 2.0: Reference chord = 11.97m: Tip chord = 1.43m: Leading edge sweep =

550: Trailing edge sweep = - 20

0.

Option (B) figure 62:-


2: Aspect

ratio = 3.29: Reference chord = 11.45m: Tip chord = 3.18m: Leading edge sweep

= 550: Trailing edge sweep inboard = -14.98

0 / outboard = 48

0.

Option (C)figure 63:-


2: Aspect

ratio = 2.48: Reference chord = 11.97m: Tip chord = 3.12m: Leading edge sweep

= 600: Trailing edge sweep = 38.7

0.

Mach Number

Ra

te o

f T

urn

Lift limit curve

Increasing lift or

reducing W/S

Increasing T/W or L/D

Speed for max instantaneous rate of turn

Speed for max sustained rate of turn

Sustained manoeuvre boundary

Structural limit (independent of aerodynamics and geometry



88

Figure 61:- Wing option (A) final initial sizing iteration.

Figure 62:- Wing option (B) final initial sizing iteration.



89

Figure 63:- Wing option (C) final initial sizing iteration.

In order to determine the aerodynamic moment about the aircrafts aerodynamic centre

for the as drawn planforms the mean aerodynamic chord for each was determined

using both the geometric method, and the calculation method (as a verification of the

geometric analysis) presented by Raymer reference 21 and shown in figure 55. The

mean aerodynamic chord c- is defined as the chord length that when multiplied by the

wing area, the dynamic pressure, and the moment coefficient about the aerodynamic

centre, yields the value of the aerodynamic moment about the aircrafts aerodynamic

centre, and can be calculated from:-

b/2

c- = 1/S ∫-b/2 c

2 dy (3.24)

Where c is the local value of chord length at any spanwise location, the spanwise

location and c- is the mean aerodynamic chord (MAC) determined by the geometric

analysis based on CATIA V5 model geometry of the wing planform options A, B, and

C, and the baseline F-35C is shown below in figures 64 through 67. As a geometry

check the numerical method was applied to each wing planform option and given

below each respective geometric analysis below.

The location and length of the mean aerodynamic chord is important when locating

the wing with respect to the fuselage, because the wing is located so that some

selected percentage of the MAC is aligned with the aircrafts centre of gravity, which

provides a first estimate of wing position to attain the required degree of stability.



90

Figure 64:- Determination of the mean aerodynamic chord for option (A).

As a verification check the mean aerodynamic chord length and location was

calculated for wing option A from: -

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 11.975 (1 + 0.11 + 0.11

2) / (1 + 0.11) = 8.067m

y- = (b/6) (1 + (2 / 1 + ) (3.26)

y- = (13.58/6) (1 + (2 x 0.11)) / (1 + 0.11) = 2.488m

The calculated values were lower than those produced by the geometric method by

17mm for c- chord length, and 13mm for y

- chord location from the aircraft

centreline and therefore in this case the geometric analysis gave a valid

approximation.

CRoot = 11.975m

CTip = 1.426m

C- = 8.084m

F-35C Outline to scale

for reference.

MAC

50%

CL

CRoot

CTip

y-= 2.501

b/2



91

For non trapezoidal wings or combinations of two trapezoidal planforms as was the

case with option (B) the approaches of splitting the wing in two was used for this

analysis of the wing mean aerodynamic chord and this is shown below, in figures 65.

Figure 65:- Determination of the mean aerodynamic chord for option (B) by

breaking the wing into Inboard (Inbd) and Outboard (Outbd) sections.

As a verification check the mean aerodynamic chord length and location for both

inboard and outboard sections were calculated for wing option B from: -

Inboard:-

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 11.445 (1 + 0.43 + 0.43

2) / (1 + 0.43) = 8.617m

y- = (b/6) (1 + (2) ) / (1 + )

(3.26)

Outbd

Inbd


for reference.

CL

C Inbd Root

CTip

C- = 8.640m

MAC Inbd

MAC Outbd

C- = 4.149m

y-= 2.640m

y-= 1.656m

b/2

CRoot = 11.445m

CTip = 4.989m

CRoot = 4.989m

C Outbd Root

50%

50%



92

y- = (7.62/6) (1 + (2 x 0.43)) / (1 + 0.43) = 1.652m

For the Inboard wing calculated the value for c- chord length was lower than those

produced by the geometric method by 23mm, and for y- chord location from the

aircraft centreline the calculated value was lower by 4mm therefore in this case the

geometric analysis gave a valid approximation.

Outboard:-

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 4989 (1 + 0.64 + 0.64

2) / (1 + 0.64) = 4.157m

y- = (b/6) (1 + (2 / (1 + )

(3.26)

y- = (11.40/6) (1 + (2 x 0.64)) / (1 + 0.64) = 2.641m

The calculated values were higher than those produced by the geometric method by


- chord location from the aircraft centreline

and therefore in this case the geometric analysis gave a valid approximation.

In order to determine the length and location of the mean aerodynamic chord for the

combined wing the outboard chord length was subtracted from the inboard chord

length to give c- = 4.460m and the values of chord location were added together y

- =

4.293m, mapping these values onto the geometry gave a c- value = 4.835m or =

375mm more than the calculated value which was attributed to the 70 difference in

sweep angle between the leading and trailing edges of the outboard wing.

Therefore for the purposes of this conceptual design study for the option B wing the

geometric value for c- and the calculated value for y

- were used for all further

analysis and empennage sizing, therefore for the combined option B wing planform:-

c- = 4.835m

y- = 4.293m



93

Figure 66:- Determination of the mean aerodynamic chord for option (C).


calculated for wing option C from: -

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 11.974 (1 + 0.26 + 0.26

2) / (1 + 0.26) = 8.411m

y- = (b/6) (1 + (2 / 1 + ) (3.26)

y- = (19.02/6) (1 + (2 x 0.26)) / (1 + 0.26) = 3.824m





CRoot

CL

50%

CTip CRoot = 11.974m

CTip = 3.120m


for reference.

MAC

C- = 8.412m

b/2

y-= 3.826



94

The values from all three wing options lead particulars were compared with the

baseline F-35C in table 6 below.

Figure 67:- Determination of the mean aerodynamic chord for F-35C.


calculated for the F-35C wing from: -

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 7.412 (1 + 0.18 + 0.18

2) / (1 + 0.18) = 5.077m

y- = (b/6) (1 + (2 / 1 + ) (3.26)

y- = (13.11/6) (1 + (2 x 0.18)) / (1 + 0.18) = 2.518m

50%

F-35C outline to 230-5 OML.

CRoot = 7.412m

CRoot

b/2

CL

CTip

CTip = 1.358m

C- = 5.081m

MAC

y-= 2.522m



95





Table 6: - Wing configuration summary.

Parameter F-35C wing Option A

wing Option B

wing Option C

wing

Span (m) 13.11 13.56 19.02 19.02

CRoot (m) 7.412 11.973 11.445 11.973

CTip (m) 1.358 2.036 3.176 3.119

0.18 0.17 0.28 0.26

Sweep angle 340 520 550 600

Area m2 57.478 94.984 109.176 143.572

Loading (kg/m2)

560 339 295 224

Aspect ratio 2.99 1.94 3.31 2.52

MAC (m) 5.077 8.179 4.835 8.411

Location of MAC (m)

2.518 2.588 4.293 3.824



96

Figure 68:- Original option (B) on FB-24 fuselage with F-35C empennage.

From the analysis conducted to this point the wing sizes in terms of span required to

meet the lower wing loading ranges in the initial sizing estimates were in the authors

view becoming too large to remain stiff enough to combat the effects of flutter and

aeroelaticity even with advanced composite aeroelastic tailoring at the speeds of

Mach 1.6 without becoming unacceptably heavy i.e. greater than 7% of MTO, for

example the option (B) wing to meet a wing loading of 290kg/m2 had a span of

19.02m (62ft), shown above in figure 68. The option (C) wing had also grown in span

and chord to achieve its target of 221kg/m2 to a size that would not permit a four tail

empennage layout on the optimum stretched fuselage and would foul the ruddervator

layout at the wing trailing edge.

The cantilever wing, because of its limited spar depth tends to be inherently flexible,

and early low speed wings were made as thick as possible, particularly at the root and

had cantilever ratios (semispan to root thickness) in the order of 10, and wings of this

low cantilever ratio were relatively stiff in bending but their torsional stiffness was

still low. By the 1940‟s cantilever ratios of the order of 15 were common and serious

consideration had to be given to aeroelastic effects in the design of fighter aircraft

wings, the option B and option C wings had a cantilever ratio of 26.8 in their original

form, therefore aeroelastic effects would be a serious issue. Sweepback also induced

additional torsion because the applied loads on the wing act significantly aft of the

wing root, so that the stiffness requirements would be greater than for a trapezoidal

wing of the same span.

Wing twisting in response to aileron deflection decreases the available rolling

moment as a function of dynamic pressure that is the speed2, because as speed

19.02m

18.29m



97

doubles the associated loads are quadrupled. In the worst case aileron reversal can

occur, this is where the control surface loads deflect the wing rather than the air.

Under these conditions the rolling moment obtained is in the opposite direction to that

commanded. This raised issues undesirable wing bending and possible reduced tip

AoA (angle of attack) as well as control reversal manoeuvring the planforms at Mach

1.6 for both planforms B and C if the low wing loading values proposed in the initial

sizing were to be achieved. Moreover to achieve the required stiffness the wing skins

would need thick and relatively heavy even in CFC, even the F/A-18A composite

wing suffered from a marked loss in roll rate in transonic flight due to its relative

flexibility, in spite of having a differential tailplane, and inset ailerons (Reference:-

25:- Fundamentals of Fighter Design 2nd

Edition: Whitford. R: Published by Airlife

Publishing 2005) although this was resolved by software re-writes. Also the

undercarriage layout would require a high degree of modification to support these

wing sizes during ground handling.

In order to address these issues at this early point in the design process a single point

wing loading was selected to enable high torsional stiffness, high sweep, high taper –

ratio, low aspect ratio versions of all three planforms to be devised as shown below.

The final wing loading selected after detailed assessment of comparable fighter /

interdictor aircraft was 388kg / m2 which reduced the wing area for all three

planforms to a standard of 83m2.

Another point of concern with respect to the option A wing was the very low taper

ratio of 0.11, which was considered as a means of reducing wing structural weight,

because as decreases from 1.0 (for a rectangular wing) to 0 (for a triangular wing)

the preponderance of the lifting force shifts inboard, closer to the wing root i.e. the

centroid of the lift distribution (c.p. centre of pressure) moves closer to the wing root.

This in turn reduces the moment arm from the root to the centre of pressure and

consequently the bending moment at the root decreases, but the lift remains the same,

and as a result the wing structure could be made lighter which would clearly benefit

the FB-24 and the A-24 aircraft. (This was supported by references 18, 21, and 22.)

However as highlighted in reference 18, as the taper ratio is decreased, the region

where flow separation first develops moves outward towards the tip, and if is

reduced to 0 the separated flow region occurs at the wing tip area, which would result

in a total loss of aileron control, this would be unacceptable for either the manned FB-

24 or the A-24 UCAV, or indeed any other aircraft. Therefore after analysing both the

F-35C and the F/A-22A which have similar control requirements to the AIA system

aircraft and have taper ratio‟s of = 0.18, and = 0.17 respectively the author

considered increasing the option A taper ratio to = 0.18 at this stage as a

precautionary measure against later redesign was incorporated into the planform A

wing maturation and reduced the impact on the amount of work required at a later

stage in the design.

The redesign was affected by reducing the forward sweep angle of the leading edge to

520 and the trailing edge to -15.32

0 from the original 55

0 and -20

0 respectively, and

this combined with cropping the tip at 900 to the leading edge achieved an increase in

tip chord to CTip = 2.030m from the original CTip = 1.430m giving a = 0.17. This



98

reconfigured wing was used for the rest of the conceptual design study, and the MAC

analysis is given below in figure 69.


calculated for the F-35C wing from: -

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 11.473 (1 + 0.18 + 0.18

2) / (1 + 0.18) = 7.859m

y- = (b/6) (1 + (2 / 1 + ) (3.26)

y- = (13.72/6) (1 + (2 x 0.18)) / (1 + 0.18) = 2.636m

The calculated value for c- was 455mm higher than those produced by the geometric

method because of the right angle tip crop, and the value for y- was 18mm higher

than the geometric analysis and the geometric values were used for this conceptual

design, see figure 69 below therefore the geometric values were used throughout.

Figure 69:- Determination of the mean aerodynamic chord for the revised wing

option A.

MAC

C- = 7.404m


CTip = 2.030m

CRoot

50%


for reference.

CL

y-

= 2.588m

b/2



99


option B.

As a verification check the mean aerodynamic chord length and location for both

inboard and outboard sections were calculated for wing option B from: -

Inboard:-

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 11.445 (1 + 0.46 + 0.46

2) / (1 + 0.46) = 8.736m

y- = (b/6) (1 + (2 / (1 + )

(3.26)

y- = (6.80/6) (1 + (2 x 0.46)) / (1 + 0.46) = 1.490m

For the Inboard wing calculated the value for c- chord length was lower than those

produced by the geometric method by 31mm, and for y- chord location from the

aircraft centreline the calculated value was lower by 4mm therefore in this case the

geometric analysis gave a valid approximation.


for reference.

CRoot = 5.351m

CTip = 5.351m

CRoot = 11.445m

CTip = 2.289m

CL

CRoot

C- = 4.025m

C- = 8.767m

MAC

MAC

y-

= 1.494m

y-

= 1.494m b/2

50%

50%

CTip



100

Outboard:-

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 5.351 (1 + 0.42 + 0.42

2) / (1 + 0.42) = 4.010m

y- = (b/6) (1 + (2 / (1 + )

(3.26)

y- = (6.92/6) (1 + (2 x 0.42)) / (1 + 0.42) = 1.494m


13.5mm for c- chord length, and 3.5mm for y

- chord location from the aircraft

centreline and therefore in this case the geometric analysis gave a valid

approximation.

In order to determine the length and location of the mean aerodynamic chord for the

combined wing the outboard chord length was subtracted from the inboard chord

length to give c- = 4.726m and the values of chord location were added together y

- =

2.984m, mapping these values onto the geometry gave a c- value = 6.098m which

was 1.372mm more than the calculated value which was attributed to the 28.50

difference in sweep angle between the leading and trailing edges of the outboard

wing.

Therefore for the purposes of this conceptual design study for the option B wing the

geometric value for c- and the calculated value for y

- were used for all further

analysis and empennage sizing, therefore for the combined option B wing planform:-

c- = 6.098m

y- = 2.984m



101


option C.


calculated for wing option C from: -

c- = (2/3) cRoot (1 + +

2) / (1 + (3.25)

c- = (2/3) 10.100 (1 + 0.20 + 0.20

2) / (1 + 0.20) = 6.958m

y- = (b/6) (1 + (2 / 1 + ) (3.26)

y- = (13.72/6) (1 + (2 x 0.20)) / (1 + 0.20) = 2.668m

The calculated values were identical to those produced by the geometric method for

c- chord length, and y

- chord location from the aircraft centreline and therefore in

this case the geometric analysis gave a valid approximation.


for reference.

50%

CRoot

CTip = 2.020m


MAC

C- = 6.958m

b/2

CL

y-

= 2.668m



102

The values from all three revised wing options lead particulars were compared with

the baseline F-35C in table 7 below.

Table 7: - Wing configuration summary.

Parameter F-35C wing Option A

wing Option B

wing Option C

wing

Span (m) 13.11 13.72 13.72 13.72

CRoot (m) 7.412 11.473 11.445 10.100

CTip (m) 1.358 2.030 2.289 2.020

0.18 0.18 0.20 0.20

Sweep angle 340 520 550 600

Area m2 57.478 83.0 83.0 83.0

Loading (kg/m2)

560 388 388 388

Aspect ratio 2.99 2.27 2.27 2.27

MAC (m) 5.077 7.404 6.098 6.958

Location of MAC (m)

2.518 2.618 2.984 2.667

From the Mean Aerodynamic Chord (MAC) data obtained graphically and by

calculation the location of the wings Centre of Pressure (C of P) or Aerodynamic

Centre (a.c.) could be calculated which in subsonic flight is 25% of the MAC, but in

supersonic flight this moves back to along the wing to up to 48% of the wings MAC.

The location of the wing relative to the aircraft centre of gravity could now be

estimated as a percentage of the MAC for a given aircraft configuration e.g. 25 to

30% for a stable aircraft with aft tail: 40% (approx) for an unstable aircraft with aft

tail and 15 to 20% for an unstable aircraft with foreplanes (reference 21).

Table 8: - Wing configuration layout.

Parameter F-35C wing

Option A wing

Option B wing

Option C wing

a. c. Mach < 1.0.

1.269m 1.851m 1.525m 1.740m



103

a. c. Mach > 1.0.

2.437m 3.554m 2.927m 3.340m

Location relative to

aircraft C of G. 2.031m 3.144m 2.439m 2.783m

¼ Chord sweep

23.90 43.90 47.60 55.20

Aircraft C of G position

8.798 10.111m 11.042 11.200

From this analysis the wings could be located on the fuselage relative to the aircraft

estimated C of G.

At this stage in the conceptual design in order to reduce fuselage drag still further the

concept of the twin tandem canopy was dropped in favour of positioning the UCAV

commander in a virtual enclosure directly behind the pilot occupying 2/3rds of the

volume devoted to the lift fan engine in the F-35B and the auxiliary fuel tank in the F-

35A and F-35C variants between the intake ducts. Ingress and egress would be via a

chevron hatch immediately above the enclosure. This had four major benefits for the

FB-24 configuration which were: - no net increase in drag from extra crew station:

original F-35 common forward hinging canopy retained: reduce risk of canopy

delamination and cracking as experienced on the larger F/A-22A canopy in SDD

phase, and greater durability in bird strike: reduced risk of EM emissions, and reduce

risk of flash blinding from special stores. Current ITAR restricted indicate that the

UCAV commander could fly both the FB-24 and the A-24 in the common virtual

environment.



104

Figure 72:- Determination of the wing position relative to the aircraft C of G for

the revised wing option A mounted on the NB1 aircraft.


the revised wing option B mounted on the NB3 aircraft.

18.29m

10.111m

CL

CL

C of G

a.c.

MAC

1.851m

¼ Chord = 43.90 sweep

MAC

18.29m


1.525m

11.042m C of G



105


the revised wing option C mounted on the NB5 aircraft.

The aerodynamic centre location for the wing configurations at subsonic speeds was

determined from: -

x ac = y- tan LE + 0.25 c

- (3.27)

And at supersonic speeds the aerodynamic location for the wing configurations was

determined from: -

x ac = y- tan LE + 0.48 c

- (3.28)

Where the leading edge of the reference wing root chord was taken as x = 0, and the

values quoted are actual length on of the MAC.

CL

MAC


18.29m

C of G 10.200m

a.c.

1.740m



106

Empennage sizing: - From this data the empennage was sized for each configuration

using the methodology detailed above in section 3.2.3.3 as follows:-

Using the total area equations 3.22 and 3.23 with the historical cHT and cVT values

giving:-

SHT = 0.476 C-W SW / LHT (3.29)

SVT = 0.070 b W SW / LVT (3.30)

Where:-

LHT = 29.9% for F-35C of fuselage length (model measurements):

LVT = 28.5% for F-35C of fuselage length (model measurements):

SW = Planform area of each option:

bW = Wing span of each option:

C-W = Wing mean aerodynamic chord for each option.

(This was determined by geometric method shown in figure 55.)

For F/A-22A the moment was approximately 30% for the vertical tails LVT and

approximately 36% for the horizontal tail LHT were as for the YF-23 the approximate

value for the ruddervators was LRV = 40% from model measurements the moment

arm. Therefore the calculated value for each planform based on an initial as draw tail

the taking the distance between the 0.25% wing chord interaction point with the MAC

and the intersection of the 0.25% tail chord intersection with the reference tail MAC,

and the same method was used for the vertical tails which gave the following values:-

Option A: - LHT = 7.316m = 40% fuselage length:

LVT = 6.350m = 35% fuselage length.

Option B: - LHT = 6.526m = 36% fuselage length:

LVT = 5.082m = 27% fuselage length.

Option C: - LRV = 7.816m = 43% fuselage length:

LHT = 5.372m = 29% fuselage length.

These values were below the range of 45 – 50% given in reference 21 but were

consistent with F-35C and F/A-22A values but below that of the value measured for

YF-23 from both model and 1/48 scale drawings supplied by the Collectaire accurate

resin model company USA, with the YF-23 kit, also supported by drawing

measurements from reference 16.



107

For the conventional four tail versions of the FB-24 and A-24 aircraft namely NB1:

NB3: and NB4 the tail sizings were as follows:-

NB1 wing option A:-

SHT = 0.4046 C-W SW / LHT (3.29)

SHT = 0.4046 (7.404 x 83) / 7.316 = 34.0m2

Volume coefficient reduced by 15% was all moving (reference 21) and area

reduced by 10% as aircraft have “active” FBL computerised control system

(reference 21).

Therefore total tail horizontal area SHT = 30.6m2

SVT = 0.070 b W SW / LVT (3.30)

SVT = 0.070 (13.72 x 83) / 6.350 = 12.6m2

Therefore total vertical tail area SVT = 12.6m2

NB3 wing option B:-

SHT = 0.4046 C-W SW / LHT (3.29)

SHT = 0.4046 (6.098 x 83) / 6.526 = 31.4m2



(reference 21).

Therefore total horizontal tail area SHT = 28.3m2

SVT = 0.070 b W SW / LVT (3.30)

SVT = 0.070 (13.72 x 83) / 5.082 = 15.7m2




108

NB5 wing option C:-

SHT = 0.4046 C-W SW / LHT (3.29)

SHT = 0.4046 (6.958 x 83) / 7.816 = 29.9m2



(reference 21).

Therefore total horizontal tail area SHT = 26.9m2

SVT = 0.070 b W SW / LVT (3.30)

SVT = 0.070 (13.72 x 83) / 5.372 = 14.8m2


The values for all three options for total HT and VT area and their ratios to the wing

area are shown in table 9, together with the individual tail areas and the ruddervator

area.

As a reality check the respective SHT / SW and SVT / SW ratio values for the three

options were compared with the respective SHT / SW and SVT / SW ratio values for

similar size four tail combat aircraft recorded in table 4.1 of reference 19.

For the V tail versions of the FB-24 / A-24 aircraft namely NB2, NB4, and NB6 the

tail sizings were as follows using the cumulative size method from reference 21 page

125:-

NB2 wing option A:-

For a V tail the HT and VT areas were estimated as above then added together to give

the same total surface area for the two ruddervators as would have been required for

two separate horizontal tails, and two separate vertical tails: -

SRV = SHTc + SVTc (3.31)

SRV = 30.6 + 12.6 = 43.2m2 – 4.3m

2 = 38.8m

2

Total area reduced by 10% as aircraft have “active” FBL computerised control

system, and HT volume coefficient reduced by 15% because the ruddervator

surfaces are all moving (reference 21).



109

NB4 wing option B:-





SRV = 28.3 + 15.7 = 44.0m2 – 4.4m

2 = 39.6m

2

Area reduced by 10% as aircraft have “active” FBL computerised control



NB6 wing option C:-





SRV = 26.9 + 14.8 = 41.7m2 – 4.2m

2 = 37.5m

2

Area reduced by 10% as aircraft have “active” FBL computerised control



Comparing the total tail and total wing area ratios calculated for the three options in

table 9 within those of the F-15: F/A-18: and MiG 25 showed that the tails were about

average for both the HT and VT for all options. These were deemed suitable for the

initial layout, but were possibly conservative when compared with the F-35C and

F/A-22A, as detailed below.

The values for F-35C reference horizontal tails were 11.017m2 per tail equal to

22.034m2 for total horizontal tail area, and the F-35C reference vertical tails had a

value of 4.547m2 per tail equal to 9.094m

2 based on model measurements of the 230-5

OML.

The values for the F/A-22A reference horizontal tails were 12.630m2 per tail equal to

25.26m2 for total horizontal tail area, and the F/A-22A reference vertical tails had a

value of 16.65m2 per tail equal to 33.300m

2 based on public domain data.



110

Table 9: - Tail sizing results for HT / VT and RV configurations.

Option.

Total tail areas and ratios to total wing areas.

HT (m2) SHT / SW VT (m2) SVT / SW RV (m2) SRV / SW

A 30.6 0.37 12.6 0.15 38.8 0.47

B 28.3 0.34 15.7 0.18 39.6 0.48

C 26.9 0.32 14.8 0.17 37.5 0.45

NB1 wing option A:-

Individual reference horizontal tail surface area = 15.3m2:

Individual reference vertical tail surface area = 6.3m2.

NB3 wing option B:-



NB5 wing option C:-



NB2 wing option A:-

Individual reference ruddervator tail surface area = 19.4m2.



111

NB4 wing option B:-


NB6 wing option C:-


F-35C baseline reference aircraft:-

These results were consistent with those from the F-35C (230-5) reference model

which had values of: -

SHT / SW = 0.38:

SVT / SW = 0.20:


Individual reference vertical tail surface area = 5.61m2:

Wing reference area = 57.5m2.

From the above empennage sizing exercise the two layouts investigated namely twin

vertical tails and twin horizontal tails, or twin ruddervators was a wise choice not only

for reasons of signature and drag reduction. There were also structural and control

effectiveness implications, this was because a tall swept fin loses a large portion of its

effectiveness at high speed as a result of aerodynamic loads twisting the tip. Typical

values (quoted in reference 25) of fin efficiency loss from rudder distortion and

spanwise twist resulting from bending were 20 to 25%. The use of twin fins can

overcome this problem of tall flexible single fins, provided that they are set far

enough apart laterally to overcome mutual biplane interference.

The main supporting argument for the implementation of twin fins or ruddervators for

the FB-24 and A-24 empennage was that there would always be one vertical surface

in relatively clean flow to provide directional stability regardless of the aircrafts AoA

/ sideslip combination. The twin fin layout was also best suited for high supersonic

speeds, when mutual interference is eliminated as the Mach lines from each fin pass

behind its neighbour, this was one of the reasons for their wide spread adoption on

high speed fighter aircraft such as the F-15, F-18, F-14, F/A-22A, F-35, and Su-27 /

30, MiG-25 / 29. Although at low Mach numbers, the twin fin layout would not be the

most efficient method of improving directional stability. Additionally as explained

above when canted outwards to match the fuselage side slope angle their signature

contribution would be considerably reduced, this was exemplified by the F/A-22A,

YF-23 and F-35 aircraft, although alternatively they could be canted inward as on the

Have Blue, and the SR-71 which could have the effect of shielding the engine exhaust

as in the case of the Sabre long range bomber conceptual design the author is working

on as a private study, outside both Cranfield and BAE Systems.



112

The full aircraft OML models resulting form the above empennage sizing are shown

below in figures 75 through 81 below with their reference sizing data.

Figure 75: - NB1:- FB-24 wing option A, illustrating complete OML model with

sized vertical and horizontal tails and final wing planform, ready for control

surface sizing.

Figure 76: - NB2:- FB-24 wing option A, illustrating complete OML model with

sized ruddervators and final wing planform, ready for control surface sizing.

Vertical tails area = 6.3m2 each including

rudder. 520 sweep leading edge and trailing

edge sweep of 650.

Spigot mounted horizontal tails area = 15.3m

2 each.

520 sweep leading edge

and 150 tailing edges.

Total reference wing area = 83m2.


Spigot mounted „Ruddervator‟ tails exposed

area = 5.500m2 each. 55

0 leading edge sweep.



113

Figure 77: - NB3:- FB-24 wing option B, illustrating complete OML model with


surface sizing.

Figure 78: - NB4:- FB-24 wing option B, illustrating complete OML model with



2 each. Leading edge

sweep 550 and 28.5

0 trailing edge

sweep.

Vertical tails area = 7.9m2

each including rudder. 550

leading edge sweep and 840

trailing edge sweep.



Spigot mounted „Ruddervator‟ tails exposed area = 5.129m2 each.

Leading edge sweep 400.



114

Figure 79: - NB5:- FB-24 wing option C, illustrating complete OML model with


surface sizing.

Figure 80: - NB6:- FB-24 wing option C, illustrating complete OML model with


Vertical tails area = 7.4m2 each including rudder.

Leading edge sweep 600 and 83

0 trailing edge

sweep.



2 each. Leading edge

sweep 600 and trailing edge

sweep of 290.

Spigot mounted „Ruddervator‟ tails exposed area = 5.445m2 each.

Leading edge sweep 550.




115

Figure 81: - F-35C baseline illustrating complete OML model with reference

sized vertical and horizontal tails and wing planform, for comparison with the

NB aircraft sizing above.

Conventional horizontal tail layouts: - The horizontal tail surface leading edges of

NB options 1, 3, and 5 shown above in figures 75, 77, and 79 respectively were

planform aligned with the wing leading edges for frontal signature reduction, as were

the trailing edges of NB 3 and NB5.

However trailing edge alignment was not possible on option 1 because this would

have resulted in a large portion of the tail overhanging the spigot with an impact on

spigot loads and on the sizing of the actuator required to drive the tail surface (current

EHA shown in figure 84 was at the limits contained within the boom IML), increasing

rear fuselage weight, attempts were made to move the surface forward into a clipped

flap as in the F-35C, and F/A-22A but to achieve a structurally safe root chord to tip

chord for the exposed surface the trailing edge spar of the wing torsion box would

have to be kinked which in the authors view was unacceptable. Therefore NB1 failed

on one count of empennage planform alignment although inverse alignment was

achieved with the trailing edge of the port horizontal tail being a sweep continuation

of the starboard wing and vice versa for the starboard horizontal tail trailing edge

which may have some RCS merit.

None of the horizontal tails were given a dihedral slope like that which has been

applied to the current F-35C in publicly released drawings, or that applied to the X-

35C Concept Demonstrator Aircraft, instead they were aligned with the wing to

reduce drag and potential radar reflections. The exposed areas of these surfaces were

consistent with F-35C and F/A-22A as well as historical data from reference 19 table

4.1, and are presented in table 10 and compared with the F-35C measurements, and

F/A-22A published values.

Vertical tails area = 5.61m2 each including rudder.

Flipper mounted horizontal

tails area = 11.02m2 each. Total reference wing area = 57.5m

2.



116

Conventional vertical fin layouts: - The vertical tails of NB options 1, 3, and 5

shown above in figures 75, 77, and 79 respectively were all canted out by 250 to the

aircraft Z, X, plane in common with the fuselage side skins to retain side aspect

planform alignment and to overcome any risk of mutual biplane interference at

subsonic speeds. This represented a decrease in the side skin cant angle from that of

the F-35C, and was a fuselage cross sectional area reduction measure to reduce wave

drag and represented the maximum reduction possible without adversely impacting on

the volume of the internal weapons bays or fuel volume and is covered in section 4 of

this thesis.

The NB1 vertical tails shown in figure 75 above were of a swept planform like that of

the F-35C, but with a leading sweep angle of 520 to match the wing leading edge

sweep angle, and a trailing edge angle of 650 which again were departures from the

sweep angles employed on the F-35C, which impacts on all major side door and

access panel chevron lines, but retains mirror commonality with the intake lips. These

changes were driven by drag reduction, exposed tail surface area given in table 10,

and structural weight considerations.

The NB3 vertical tails shown in figure 77 above were of a trapezoidal planform

similar to that of the F/A-22A, but with a leading edge sweep of 550 to match the

leading edge sweep angle, and a trailing edge sweep angle of 840 this latter angle was

driven by the need to keep the exposed tail volume on the aft fuselage booms with

adequate volume in the boom itself below the rudder to house the rudder EHA

actuator system (shown in figure 84). The trapezoidal planform was selected for NB3

vertical tails to reduce the risk of flutter which could have occurred if the larger tail

area of NB3 had used a swept surface planform similar to NB1.

The large vertical tail area calculated for NB5 shown in figure 79 raised the same

issues as those raised for NB3 which resulted in adoption of the trapezoidal planform

with a leading edge sweep of 600 common to the leading edge sweep angle, and a

trailing edge sweep angle of 830 to meet rudder actuator housing requirements.

Ruddervator layouts: - All of the ruddervator layouts NB2, NB4, and NB6 shown in

figures 76, 78 and respectively were of a similar planform to that employed on the

YF-23 (shown in figure 59 above), in having a continuous leading edge sweep and

chevron trailing edge, of two sweep angles, a lower section positive angle and an

upper section negative sweep angle. All of the ruddervators were canted outwards at

500 to the aircraft Z, X, plane to give the greatest pitch control authority and to avoid

acute corners or right angles in side elevation or front view. These all-flying

ruddervator tails were to be rotated about a single spigot by the ruddervator EHA

actuators, and could be driven differentially for yaw and roll control. These canted

surfaces because of their location on the aft fuselage booms would also act as shield

for the engine exhaust in all angles except immediately above, behind or below the

aircraft. On layouts NB2, NB4, and NB6, the wing trailing edge controls also provide

roll control and lift augmentation, and in combination with the ruddervators function

as speed brakes and rudders. For straight line deceleration, the FCS commands the

outer ailerons to deflect up and the inboard flaps to deflect down, thus producing a

decelerating force but creating no other moments.



117

Yaw control would be provided by doing this on one side only. This augmentation

permitted a reduction in the exposed area of the ruddervator control below that which

would be required if the ruddervators had no augmentation, for example the exposed

ruddervator area for the YF-23 which employed the same wing control surface

augmentation to that described above was 4.739m2 for each ruddervator and the YF-

23 had a reference wing area of 87m2 according to public domain data (Reference 26:

- Advanced Tactical Fighter to F-22 Raptor Origins of the 21st Century Air

Dominance Fighter: by Aronstein D. C., Hirschberg M. J., and Piccirillo A. C.:

Published by American Institution of Aeronautics and Astronautics 1998). If no

augmentation had been used from the cumulative area calculations based on reference

21 methodology the ruddervator total reference area would have been 25m2 resulting

in a larger exposed area. Similar wing control surface augmentation was used on the

F-117A to reduce ruddervator tail size, (Reference 27): - Have Blue and The F117A

evolution of the “Stealth Fighter”: by Aronstein D. C. and Piccirillo A. C.: Published

by American Institute of Aeronautics and Astronautics 1997). Subsequent assessment

of straight line deceleration and positive yaw control indicated that the effect of the

ailerons could be incorporated into single unit flaperons and toeing in the ruddervators

(i.e. both ruddervators turning in towards the centreline) in combination with

downward deflection of the single unit flaperons could produce the same breaking

effect, therefore the need for the ailerons for the NB1 and NB2 configurations was

questioned, as described below.

The NB2 ruddervator geometry was driven by the need to reduce to a minimum

control surface overhang on the aft fuselage boom in order to position the diving

spigot as close to the centre of the ruddervator as possible to reduce the loads on the

spigot and the driving forces required from the actuator, (with the corresponding

actuator size requirement shown in figure 82). In order to achieve this the leading

edge sweep angle of 550 was selected along with a trailing edge sweep of -48

0 for the

top section and 260 for the lower section imparting chevron trailing similar to that of

the YF-23 ruddervators, and although alignment with F-35C features was not

maintained for the new build aircraft all the features would be changed to match the

ruddervator geometry. The exposed planform area was 5.5m2 for each tail, which was

considered adequate with the previously mentioned augmentation.

The same drivers were applied to the ruddervators for NB4, which resulted in a

leading edge sweep of 400 and a trailing edge chevron with a -52

0 sweep for the upper

section and a 270 for the lower section. The exposed planform area for this surface

was 5.1m2 for each tail.

For NB6 the same planform as NB2 was adopted to elevate the issues discussed above

in terms of actuators and overhang, and this geometry was modified to give an

exposed planform area of 5.4m2 for each tail surface.

In order to prevent clashes between the ruddervators and the wing trailing edge

control surfaces a clearance separation of 0.4m aft of the reference wing trailing edge

was established to permit free rotation of the ruddervators and the trailing edge flaps,

without the need for flap trailing edge cutouts as used on the F/A-22A and F-35C.



118

Figure 82: - EHA actuator for horizontal tail and ruddervator actuation. Source

authors private collection.

Figure 83: - EHA actuator for rudders actuation on conventional vertical tails,

and trailing edge flap and aileron actuation on all aircraft configurations. Source

authors private collection.

Length = 6.94m.

Height = 3.45m.

Height = 1.12m.

Length = 2.75m.

Weight = 116kg.

Weight = 32kg.



119

Table 10: - Exposed tail areas for all configurations.

Configuration Individual

Horizontal tail area in m2.

Individual Vertical tail area

in m2.

Individual Ruddervator tail

area in m2.

NB1 6.7 5.8 N/A

NB2 N/A N/A 5.5

NB3 5.8 7.3 N/A

NB4 N/A N/A 5.1

NB5 5.7 6.6 N/A

NB6 N/A N/A 5.4

F-35C (ref) 5.8 4.6 N/A

Control surface and high lift device sizing: - The next stage before full aircraft drag

polar analysis was to size the wing control surfaces (and the rudder for the

conventional four tail aircraft layouts), which was addressed below, using the F-35C

sizes to determine a wing area to control surface area relationship, as an initial starting

point for analysis. The following control surfaces were sized in this manner: - rudder

based on tail area: trailing edge flaps based on wing area: ailerons based on wing area:

and finally the leading edge flap. The results in terms of the ratio of control surface

area to wing reference area were as follows for the F-35C: -

Total trailing edge flap area to reference wing area = 0.10 = 5.76m2:

Total aileron area to reference wing area ratio = 0.05 = 2.70m2:

Total leading edge flap to reference wing area ratio = 0.08 = 4.672m2:

Total rudder area to total vertical tail reference area ratio = 0.24 = 2.724m2.

The final sizing of these control surfaces would be based upon dynamic analysis of

control effectiveness, including structural bending and control system effects,

however these ratios taken from a real aircraft of similar type and compared with data

from table 4.1 in reference 19 were considered by the author as being adequate for

this initial concept design study with constitutes a proposal submission.



120

The reference wing area of the F-35C 230-5 OML represented 69.25% of the

reference wing area of all Advanced Interdiction Aircraft configurations i.e. 57.478m2

for the F-35C compared to 83m2, therefore for the purposes of this study the sizes of

the reference F-35C control surfaces actual sizes should have increased by a factor of

0.3075 = 30.75% to yield the same effectiveness on the larger AIA wing. However

when the above ratios were applied the sizes were larger than predicted as follows: -

Ailerons = 83 x 0.05 = 4.15 = 2.075m2 (per aileron):

Trailing edge flaps = 83 x 0.10 = 8.30 = 4.150m2 (per trailing edge flap):

Leading edge flap = 83 x 0.08 = 6.64 = 3.32m2 (per leading edge flap).

For the vertical tailed configurations rudder size determination used the same scaling

ratio method applied to the exposed vertical tail area each four tail configuration as

follows:-

NB1 rudder area = 5.8 x 0.29 = 1.682m2 (per rudder):

NB3 rudder area = 7.3 x 0.29 = 2.117m2 (per rudder):

NB5 rudder area = 6.6 x 0.29 = 1.914m2 (per rudder).

From these results the control surfaces were modelled onto the exposed wings of

surface models of each configuration with the leading edge flaps occupying the

leading edges from root to tip and the trailing edge flaps occupying the inboard 50%

of the wing trailing edge, and the ailerons occupying the outboard 50% of the wing

trailing edge, as shown in figures 85 to 90. This enabled location of the leading and

trailing edge spars for each configuration and defined the outer boundaries of the

wing torsion box. The ailerons and trailing edge flaps were to be driven by a single

MOOG EHA actuator each, of the type shown in figure 83. Were as the leading edge

flaps were driven by four segmented MOOG rotary actuator of the type shown below

in figure 84.

The rudders were driven by same type of EHA as used for the wing trailing edge

control surfaces. The rudders began at 10% of the exposed trailing edge of the vertical

tails for configurations NB3 and NB5 to clear the horizontal tail spigot bulge in the aft

fuselage booms, and extended to the tip of the tail. Because of the twin tail layout loss

of rudder efficiency at high speeds with this rudder span was not considered an

important issue for reasons explained above, as it would have been on a single tail

aircraft probably would have required an all moving vertical tail like the TSR-2, or

Vigilante. The NB1 rudders began at 5% of the trailing edge chord to prevent OML

clash and extended over the full span of the tail.



121

Figure 84: - The multi segment rotary actuator selected to drive the leading edge

flaps with four actuators located at ¼ span intervals along the leading edge, for

even load distribution. (Note assembly represents a single actuator reference

MOOG).

Figure 85: - Flight control and high lift device integration on the NB1/NB2 wing

planform illustrating (a) F-35C based layout and (b) Revision for FB-24 / A-24.

T/E Flap

L/E Flap

Tip Aileron

L/E Flap Flaperon

(a) Three surfaces.

(b) Two surfaces.



122

During design maturation studies of the „Option A‟ wing control surfaces as sized

above the aileron could only be accommodated as a spigot mounted moving wing tip

device as a result of the large size of the trailing edge flap, on the relatively short span

exposed wing as shown above in figure 85(a). This lead to concerns about: - the

effectiveness of the aileron as an aerodynamic control: the difficulty of integrating an

actuator: and structural weight penalty of reinforcing the torsion box to withstand high

torque at the wing tip rib and spar tips resulting from the ailerons actuation. The

trailing edge flap was judged by the author to extend for enough outboard and to be of

sufficient surface area to perform both the functions of a flap and an aileron for the

„Option A‟ thereby being a flaperon with the same area ratio to the reference wing as

the flaperons of the F-35A, F-35B and F-16 aircraft and located over the same extent

of the trailing edge exposed span, as shown in figure 85(b). Therefore the aileron for

this wing option was dropped for both the NB1 and NB2 configurations as shown in

figures 86 and 87. This had the benefits of reducing the weigh of the wing both in

terms of structural weight and systems weight (through elimination of two wing

actuators), although this had to be traded against provision of two small fuselage

mounted ventral airbrakes for straight line deceleration, on landing which had

individual areas of one quarter of that for an aileron as sized above, and their

actuators, for the NB1 configuration.

Figure 86: - FB-24 NB1 Configuration CATIA V5 solid model control surface

layout, illustrating the size and location of Flaperon control surfaces and

Leading edge flap high lift device in the cruise condition i.e. not deflected for

manoeuvre. The high swept wing angle to keep within the Mach wave cone

reduced the trailing edge available necessitating the adoption of Flaperons.

These surfaces are crudely sized using comparative area assessment and require

a full – scale analysis as described above which was recommended as further

work.

L/E Flap area = 3.515m2 each

T/E Flaperon area = 4.150m2 each

Rudder area = 1.682m2 each



123


layout, illustrating the size and location of control surfaces and high lift devices

in the cruise condition i.e. not deflected for manoeuvre. Wing control surfaces

and high lift devices are identical to NB1. These surfaces are crudely sized using

comparative area and require a full – scale analysis as described above which

was recommended as further work.

The proposed Control Surface Authority of these devices was estimated based on

similar legacy aircraft to be as follows: -

L/E Flaps: - 3.00 Up / 40.0

0 Down (Normal to Hinge Line):

T/E Flaperons: - 350 Up / 35


Horizontal Tails (NB1): - 300 Up / 30

0 Down (Normal to Spigot Centre Line):

Rudders (NB1): - 300 Inboard / 30

0 Outboard (Normal to Hinge Line):

Ruddervators (NB2): - 250 Inboard (limited by structure) / 30

0 Outboard

(Normal to Spigot Centre Line).


T/E Flaperon area = 4.150m2 each „Ruddervator‟ tails exposed

area = 5.500m2 each.



124



in the cruise condition i.e. not deflected for manoeuvre. These surfaces are

crudely sized using comparative area and require a full – scale analysis as

described above which was recommended as further work.

As shown in figures 88 and 89 the NB3 and NB4 configurations high wing leading

edge sweep of 550 reduced the effective trailing edge span available for the larger

control surfaces sized above, as was the case for configurations NB1 and NB2 and

this was further compounded by the novel cranked arrow wing planform of the wing,

which constrained the flap to the inboard wing. This resulted in a trailing edge flap

area only 58% that of the estimated value based on the ratio of flap size to wing

reference area. However the aileron area was increased by 13% over target and the

leading edge flap area was increased 10% to compensate in rotation and landing, and

these size increases were judged by the author as being the maximum values that the

outboard wing structure could withstand without large structural weight penalties.

The larger than estimated aileron area was also used to compensate for the fact that

the rudders were 1% smaller than estimated the target value due to the vertical tail

geometry and structural requirements of flutter as well as the systems to be mounted

on the tip rib under the tip rib fairing.

The NB3 configuration would in the view of the author require the addition of ventral

airbrakes as in the case of the NB1 configuration for straight line deceleration, where

as the NB4 configuration would employ the same breaking technique as the YF-23

and NB2 configuration described above, and would therefore not require the addition

of airbrakes.


Aileron area = 2.376m2 each

T/E Flap area = 2.444m2 each




125








L/E Flaps: - 3.00 Up / 40.0




Aileron s: - 250 Up / 25







0 Outboard





Ruddervator tails exposed area = 5.129m2 each



126






Once again the penalty for a high leading edge sweep angle required to reduce wave

drag was a reduction in trailing edge span available for high lift and flight control

surface, and the 600 sweep of configurations NB5 and NB6 was the most severe. Also

attempting to extend these surfaces chordwise to recover the estimated incurred

structural weight penalties due to the resulting reduction in torsion box chord leading

to the torsion and bending loads being carried by fewer structural members which in

turn would need to be increase in section to retain the same overall stiffness, and this

would increase the wings structural weight. After much iteration the layout shown in

figures 90 and 91 was developed which in the authors view was the most practical

compromise between, control and high lift device requirements and the stiffness and

fuel capacity requirements for the wing torsion box.

In this layout the ailerons were modelled at 77% of the estimated target value, and the

trailing edge flaps were modelled at 81% of the estimated target values, although

some small redress was achieved by increasing the leading edge flap area to 109%,

and increasing the rudder area to 102% of the estimated target values. However

longitudinal stability a key requirement in a bombing platform could still be an issue

and requires further investigation.

In the authors opinion both NB5 and NB6 would require ventral air breaks which in

addition to simultaneous deployment for straight line deceleration, would be deployed

differentially to aid the ailerons in roll control.







127








L/E Flaps: - 3.00 Up / 40.0




Aileron s: - 250 Up / 25







0 Outboard





Ruddervators exposed

area = 5.445m2 each



128

The author had reservations with both the option B and C wing control surfaces

abilities to counter any „Mach Tuck‟ (nose down attitude at transonic acceleration,

and extreme nose up attitude on deceleration which is a phenomenon high wing

sweep short wing span long fuselage configurations reference 25) which may occur

on acceleration to penetration speed, and deceleration on exiting the threat area.

Detailed sizing analysis was required to establish the capabilities of the wing and trail

control surfaces for theses configurations which could not be undertaken within the

time scale of this thesis.

Table 10 below brings together all of the control and high lift device modelled areas

for all configurations and compares this with the estimated target values.

Table 11: - Installed control surface sizings single component areas.

Control Surface.

Wing Option A (NB1 and NB2)

in m2

Wing Option B (NB3 and NB4)

in m2

Wing Option C (NB5 and NB6)

in m2

Target areas in m2

Aileron N/A 2.376 1.617 2.075

L/E Flap 3.515 3.673 3.651 3.320

T/E Flap 4.150 2.444 3.388 4.150

Rudder NB1 1.682 N/A N/A 1.682

Rudder NB3 N/A 2.091 N/A 2.117

RudderNB5 N/A N/A 1.950 1.914

From table 10 and the above it is clear that the high wing sweep and root chord

constraints had adverse effects on the trailing edge areas of each wing planform which

could be devoted to these surfaces and high lift devices. However the target areas

could be met or exceeded for the leading edge flap high lift device for all planforms.

The rudder areas of the three conventional tail options met the target for NB1, and

exceeded the target for NB5, but NB3 was just below target.

As stated above the validity of these target values needs to be established by dynamic

analysis of control effectiveness, including structural bending and control system

effects which would form part of a further more in depth study than undertaken in this

thesis, and this concluded the configuration OML sizing.



129

3.2.5: - Common FB-24 and A-24 supersonic drag analysis for Jet306 analysis

selection.

For the specified mission both manned and unmanned configurations of the Advanced

Interdiction Aircraft would spend a large proportion of flight time at supersonic

speed, therefore it was important that the aerodynamic design concentrated on

reduction of the wave drag of the aircraft.

The initial aerodynamic estimation concerned the prediction of aircraft drag and lift

and for the three configurations investigated the main focus of drag was on the

supersonic wave drag (CDw) estimation. Using the supersonic drag analysis

methodology from reference 22 pages 132-133 and Swet values obtained from

measurements taken from the CATIA V5 surface models, the values for MCrit, CDw,

and k1, for the NB1, through NB6 configurations were determined for MCrit, M = 1.4,

and M = 1.6.

Supersonic drag analysis for NB1 / NB2:-

Step 1: - MCrit, from equations 3.14 and 3.17:-

MCrit (unswept) = 1.0 – 0.065[100(t max /c)] 0.6

= 1.0 – 0.065(5.9)0.6

= 0.82

MCrit = 1.0 – cos0.6

0.25c (1.0 - MCrit (unswept)) = 1.0 – cos0.6

43.90 (1 – 0.82) = 0.852

MCDo max = 1 / (cos0.2

LE) = 1 / (cos0.2

520) = 1.10

Step 2: - CDw, from supersonic zero lift drag equation 3.32:-

CDw = 4.5 / S (Amax / l) 2 EWD (0.74+0.37 cos LE) [1-0.3M-MCDo] 3.32

Where Amax = 4.47m2:

l = 18.29m

LE = 520

S = 83m2

MCDo = 1.10

EWD = 1.4



130

CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.10-1.10] = 0.01424

CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.40-1.10] = 0.01190

CDw=4.5 /83 (4.47/18.29)2 EWD (1.4) [1.0 – 0.31.60-1.10] = 0.01122

Step 3: - k1 from the supersonic drag due to lift equation 3.33:

k1 = [AR (M2 – 1) / (4ARM

2 – 1) – 2] cos LE 3.33

k1 = [2.27(1.102 – 1) / (4 x 2.27 (1.10

2 – 1) – 2] cos 52

0 = 0.13581

k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40

2 – 1) – 2] cos 52

0 = 0.13701

k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60

2 – 1) – 2] cos 52

0 = 0.23700




= 1.0 – 0.065(5.9)0.6

= 0.82



47.60 (1 – 0.82) = 0.823


LE) = 1 / (cos0.2

550) = 1.00




l = 18.29m

LE = 550



131

S = 83m2

MCDo = 1.10

EWD = 1.1

CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.00-1.00] = 0.01230

CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.40-1.00] = 0.00906

CDw=4.5 /83 (4.47/18.29)2 EWD (1.1) [1.0 – 0.31.60-1.00] = 0.00859


k1 = [AR (M2 – 1) / (4ARM

2 – 1) – 2] cos LE 3.33

k1 = [2.27(1.002 – 1) / (4 x 2.27 (1.00

2 – 1) – 2] cos 55

0 = 0.00000

k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40

2 – 1) – 2] cos 55

0 = 0.18124

k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60

2 – 1) – 2] cos 55

0 = 0.21745




= 1.0 – 0.065(5.9)0.6

= 0.82



55.20 (1 – 0.82) = 0.821


LE) = 1 / (cos0.2

600) = 1.04





132


l = 18.29m

LE = 600

S = 83m2

MCDo = 1.10

EWD = 1.2

CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.04-1.00] = 0.01148

CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.40-1.00] = 0.00989

CDw=4.5 /83 (4.47/18.29)2 EWD (1.2) [1.0 – 0.31.60-1.00] = 0.00937


k1 = [AR (M2 – 1) / (4ARM

2 – 1) – 2] cos LE 3.33

k1 = [2.27(1.042 – 1) / (4 x 2.27 (1.04

2 – 1) – 2] cos 60

0 = 0.02689

k1 = [2.27(1.402 – 1) / (4 x 2.27 (1.40

2 – 1) – 2] cos 60

0 = 0.16235

k1 = [2.27(1.602 – 1) / (4 x 2.27 (1.60

2 – 1) – 2] cos 60

0 = 0.18955

The wave drag analysis above was inconclusive and in order to focus on a single

configuration for further analysis other factors than supersonic drag were considered

in order to down select the PWSC configuration without extensive analysis of all

three configurations; these were usable fuel volume (based on the fact that the

fuselage volumes were almost identical the wing volume became the deciding volume

factor), predicted structural weight based on an initial layout study and the predicted

number of structural members.

Although NB3 and 4 had the best predicted wave drag at M 1.6 the wing had a

fuel tank volume of 6.39m3 compared with 8.2m

3 for NB1 and NB2

configurations.



133

The NB5 and NB6 configurations although marginally better wave drag at M1.6

than the NB1 and NB2 configurations they had the smallest wing fuel tank

volume 5.63m3 compared to 8.2m

3 for NB1 and NB2 configurations.

This was considered a major risk if the NB3 / 4 or NB5 / 6 configurations were

required for endurance rather than high speed missions resulting from a mission

profile change.

The exposed wing of the of NB3 and NB4 configurations were marginally smaller

than the NB1 and NB2 configurations at 18.122m2 compared to 18.628m

2 which

would offer some degree of benefit in parasitic drag reduction, but the cracked arrow

layout would not offer the same even load distribution which should be possible with

the trapezoidal wing of the latter configuration. Loads would need to be concentrated

into a small number of relatively heavy members to obtain a high degree of structural

stiffness. Also the need to back – up the aileron control surface attachment lugs with

ribs, and house the another actuator in a relatively thin portion of the wing would

further increase structural weight.

The exposed wing area of configurations of NB5 and NB6 were larger than those of

configurations at 19.425m2 compared with 18.628m

2 for the NB1 and NB2

configurations, which would have had an adverse affect on parasitic drag values

compared with the latter configurations. The structure would also need to be heaver

outboard to maintain structural stiffness in view of the wings high sweep angle the

lower taper ratio, and the influences of the aileron loads on the wing structure, as well

as the incorporation of the actuator in the outboard wing. Also lateral stability could

be an issue with the NB5 and NB6 configurations requiring a re-examination of the

empennage design and sizing.

Also all four of these configurations would have a higher approach and landing speed

than configurations NB1 and NB2 in view of their higher wing sweep, therefore the

landing field requirements may be harder to achieve.

On balance and in view of the workload to investigate all three configurations to the

fullest extent the author selected the NB1 and NB2 configurations as the best

candidates for further study in view of being the most versatile configurations with

the greatest potential for role change and possible systems weight growth, even if the

supercruise capability was diminished, the overall requirements were versatility range

and endurance capabilities, rather than point design.



134

3.3 Configuration optimisation by parametric analysis.

3.3.1 Analysis methodology:

This design study was a detailed analysis of the down selected NB1 and NB2

configurations using analysis models created for Jet306. The following Conceptual

design technologies available within the AeroDYNAMIC V3.0 tool set were used for

the configuration analysis:-

Customer Focus – Needs, House of quality.

Design Synthesis – Aircraft configuration modeling.

Geometric Modeling – Areas and Volumes.

Aero Analysis – Parametric aerodynamic analysis.

Propulsion Modeling – Parametric based on F-135 / F-136 published data.

Constraint Analysis – Design Point.

Mission Analysis – Better Mission Fuel Fraction.

Structural / Weight Prediction – Weights analysis.

Sizing – Sized Wing Area, WTO, and TSL.

Performance Analysis - Ps

Cost Analysis – Acquisition and operating costs, life cycle.

Sensitivity / Optimization – “Best” Design, Cost trade.

With the configuration generation stage complete and initial down selection through

drag polar analysis and range analysis using the methodologies in pages 126 to 133 of

reference 22, (as detailed in appendices E) which were used to compare the six FB-

24 manned (worst case with canopy drag) configurations against the F-35C, the

detailed analysis of the two variants of the PWSC wing option was undertaken: -

1. This started with the construction and of the NB1 and NB2 aircraft

configuration analytical models in Jet306 from measurements taken from the

detailed OML models constructed in the configuration generation stage. These

models were crude by comparison with the CATIA V5 models due to the

limits of the Jet306 geometry generation tool, for example the aft fuselage

booms could not be modelled as separate items on the aircraft and had to be

incorporated as extensions to the empennage aft of the engine nozzle. Also the

intakes could not be modelled as chevrons in plan view but only swept, as in

the side elevation, and aircraft with very extensive compound curvature would

be difficult the model with this tool. The Jet306 exhibited numerous run time

errors in operation and a replacement was going to take three months to arrive

and the author is still awaiting delivery. So for the analysis in this study the

older and geometrically less sophisticated AeroDYNAMIC V2.08 was used.

2. These configurations were then analysed using the AeroDYNAMIC toolset,

and compared with the baseline aircraft and the weight targets given above.



135

3. The results of these studies were used to determine the best final basepoint

configuration, and this basepoint configuration was developed into a higher

fidelity CATIA V5 solid model for structural layout and analysis.

AeroDYNAMIC 2.08 was also a Whole Aircraft Analysis tool capable of determining

and reporting the following aerodynamic behaviour of a given configuration: - Lift:

Parasite Drag: Induced Drag: Supersonic Wave Drag: Mcrit: supporting the following

analysis: - constraint analysis: manoeuvre analysis: performance and specific excess

power analysis: stability and control analysis: sizing: weight prediction: optimisation:

and cost analysis. The above are based on geometry, mission and engine data as

detailed below. The initial data required in order to construct the Jet306

AeroDYNAMIC analysis model in the Main spreadsheet consist of data parameters

defining the fuselage, and parameters defining the wing, and empennage geometry for

determining the initial wetted area the configuration, and the fuselage is then further

defined in the geometry spreadsheet with the 20 specific geometry data points at 20

frame stations, and this much more accurate data is used for the Swet.

The core fuselage design is broken into three cylindrical sections (note cross –

sectional shapes of the items shown as cylinders and half cylinders in figure 92 may

be circular elliptical, rectangular or any other shape) in Jet306 each frame is defined

by specifying the length, midline width and centreline height of each section with the

exception that the dimensions where two adjacent sections connect must match. If the

height and width of a section are equal, the section will have a circular cross section.

If the height and width of a section differ, the section will have an elliptical cross

section. If the height and width of a section are zero, the section will effectively

become a cone.

Figure 92: - Fuselage Design Breakdown for 3 – section fuselage.

3.3.2:- NB1 and NB2 Analysis model construction.

Form the NB1 and NB2 surface models created above and the as drawn engine based

on YF-120 dimensions the input data was obtained as shown in figure 93 and 94.

Then section cuts were taken through the CATIA surface model at 20 equal distance

stations and breaking the OML into outline contour cuts, from which the width and

height for each section was measured. This data was then entered into the

AeroDYNAMIC 2.08 geometry data spreadsheet to create a basic definition of the

Fwd Mid Aft h1

w1

h1

w1 w2

h2

h2 w2

Fwd len Mid len Aft len

Cone # 1 Cylinder # 2 Cylinder # 1



136

fuselage. The airframe centre line was determined to be on the engine thrust line and

all point co – ordinates were measured relative to this, at each frame station.

Figure 93: - NB1 Section cuts for Jet306 analysis (note what appear to be surface

blemishes in the cockpit are white reflections as capture used show white as

black feature to show the cuts).

Figure 94: - Analysis engine for Jet306 data from YF-120 data used for NB1 and

NB2 evaluation constructed in CATIA V5 by the author.

This data was used with the location and size of the crew stations, fuel weight, engine

weight and C of G location, control surface sizing and location, undercarriage layout,

payload weight and volume, fuel weight and volume, to evaluate the NB1 and NB2

configurations for which the results are presented below.

Airframe mate joints denoted by orange planes.



137

Chart 1:- (a) NB1 Total predicted drag variation with Mach number: (b) NB1

Drag polar to Mach 0.85

Figure 95:-NB1 Cross-sectional area distribution which showed a marked

similarity to that of the F/A-22A as shown in figure 24.

(a)

(b)

Wave drag

Profile drag



138

Chart 2:- NB1 Lift over drag verses lift coefficient at Mach 0.85 and Mach 1.5

curves: - (a) L/D –v- CL at Mach 0.85, and (b) L/D – v – CL at Mach 1.5.

Chart 3:- NB1 Lift curve CL v

(a)

(b)



139

Chart 4:- NB1 CD – v – CL at Mach 0.85 (a) and Mach 1.5 (b).

Chart 5:-NB1 Thrust and Drag –v-Mach number.

(a)

(b)



140

Chart 6:- (a) NB1 Vn Diagram: (b) NB1 Manoeuvre Diagram.

Chart 7:- NB1 Specific Excess Power over a range of Mach numbers and

altitudes.

(a)

(b)

Stall - Limit

q - Limit

NB-1



141

Chart 8:- (a) NB2 Total predicted drag variation with Mach number: (b) NB1

Drag polar to Mach 0.85

Figure 96:-NB2 Cross-sectional area distribution which showed a marked

similarity to that of the F/A-22A as shown in figure 24.

(a)

(b)

Wave drag

Profile drag



142

Chart 9:- NB2 Lift over drag verses lift coefficient at Mach 0.85 and Mach 1.5

curves: - (a) L/D –v- CL at Mach 0.85, and (b) L/D – v – CL at Mach 1.5.

Chart 10:- NB2 Lift curve CL v

(a)

(b)



143

Chart 11:- NB2 CD – v – CL at Mach 0.85 (a) and Mach 1.5 (b).

Chart 12:-NB2 Thrust and Drag –v-Mach number.

(a)

(b)



144

Chart 13:- (a) NB2 Vn Diagram: (b) NB1 Manoeuvre Diagram.

Chart 14:- NB2 Specific Excess Power over a range of Mach numbers and

altitudes.

(a)

(b)

Stall - Limit

q - Limit

NB-2



145

Table 12:- Comparative component weights: GTOW / Landing Weight: Costs

for NB1 and NB2 based on the RAND Corporations DARCO IV model in

AeroDYNAMIC V2.08.

COMPONENT NB1:-kg / (lb) NB2:-kg / (lb)

Fuselage 2,837.6 / (6,255.7) 2,837.6 / (6,255.7)

Verticals 333.6 / (735.4) 323.2 / (712.6)

Wings 1,232.4 / (2,716.9) 1,232.4 / (2,716.9)

Undercarriage 693.1 / (1528.0) 590.4 / (1301.7)

Horizontals 753.4 / (1661.0) N/A

GROSS WEIGHTS NB1 NB2

GTOW 23,182.8 / (51109.4) 22,327.4 / (49223.5)

Landing Wt 10% fuel 12,063.45 / (26,595.4) 11,208.0 / (24,709.5)

x CofG TO 5.334m / 17.50ft 4.898m / 16.07ft

x CofG Landing 10.250m / 33.62ft 9.754m / 32.00ft

COSTS* NB1 NB2

Fly Away Cost 1st Estimate

$51,109,360 $49,223,450

Total fly away costs $27,745,394 $25,967,104

Total Q & M costs $19,856,443 $19,856,443

Total Life Cycle costs $47,601,837 $45,823,547

Stability NB1 NB2

MAC 25.21123ft 25.21123ft

xMAC 22.69141ft 22.69141ft

yMAC 7.99775 7.99775

NPSub 0.283562 0.160700

NPSup 0.517118 0.407723

SMTO Sub 0.723202 0.670459

SMlnd Sub -0.15007 -0.20884

*N.B.:- Costs are based on production run of 500 aircraft with 3 flight test

aircraft.

Based on the above values NB2 had better weight and cost values than NB1

therefore NB2 went forward to the structural layout phase detailed in section 4.



146

For the FB-24 configuration the maximum endurance is achieved at the speed for

(L/D)max which can be determined by first calculating the required value of CL, then

solving for the speed required to achieve L=W at that CL therefore for FB-24:-

CL = CDo / k = 0.014361 / 0.167051 =0.293

L = W = CLqS, q = W / CLS = 53,187.50 / 0.293 (893ft2) = 203.27lb/ft

2

And using at 34,400ft = 0.000767 slug/ft3 obtained from standard atmospheric

tables (reference 22) and the definition of q,

V = 2q / = 2 x (203.271lb/ft2) / 0.000767 slug/ft

3 = 728.05ft/s

for maximum endurance. Note that this is only the initial velocity for maximum

endurance and that as fuel is burned the velocity for best endurance will decrease. In

order to calculate the maximum endurance time, it is first necessary to determine the

magnitude of (L/D) max using equation 3.34: -

(L/D) max = (CL / CD) max = 1 / 2 k CDo = 1 / 2 0.167051(0.014361) = 10.2

The TSFC is also predicted using equation 3.35 with a = 977.5ft/s at 34,400ft and

aSL = 1116.4ft/s obtained from the standard atmospheric table (reference 22):-

ct = cSL (a / aSL) = (0.89[(lb/hr)/lbf])(977.5ft/s / 1116.4ft/s) = 0.78[(lb/hr)/lbf]

Then the endurance can be calculated using equation 3.36 with W1 = 49,223.5 and

that for W2

W2 = W1 – Wf = 53,187.46lb – 21,126lb = 31,061.46lb

E = 1/ ct CL / CD ln (W1 / W2) = 1 / 0.78 (10.2) ln (53,187.46 / 31,061.46) = 7.03h

Therefore for FB-24 the maximum endurance at 34,400ft at Best Cruise Mach

BCM is 7.33hours.

Similarly the value for maximum range is obtained by solving equation 3.36 for CL

and equation 3.37 for q:-

CDo = 3k CL2, CL = CDo / 3k = 0.014361 / 3(0.167051) = 0.239

q = W / CLS = 53,187.46 / 0.239(893ft2) = 249.20lb/ft

2



147

V = 2q / = 2 x (249.20lb/ft2) / 0.000767 slug/ft

3 = 806.11ft/s

for maximum range. As with the velocity for maximum endurance, the velocity for

best range will decrease as fuel is burned. The vale calculated for CL is now used to

calculate CD after which the maximum range is predicted using equation 3.38.

CD = 0.014361 + 0.167051 CL2 = 0.014361 + 0.167051(0.239)

2 = 0.0239

R = 2/S / ct CL2 / CD (W1

1/2 – W2

1/2)

R= 2 / (0.000767 slug/ft3) (893ft

2) 2 / 0.78 (0.239)

1/2 / 0.0239

x ((53,187.46)1/2

-(32,061.46lb)1/2

)

R = 15.315 (ft2 / lbs

2) (2.564h)(20.45)(10.563lb

1/2) = 2,167.49

Therefore for the range in nautical miles divide through by 1.69(ft/s/kn)

Therefore the FB-24 would have a range value of R = 1,282.54nmiles.

As stated above the drag values from AeroDYNAMIC were pessimistic being based

on the a oval cross section throughout the length of the fuselage and the booms having

to be treated as complete sections of the fuselage, both of theses factors have lead to

an overly pessimistic parasitic and wave drag values. The A-24 has a larger fuel

capacity than the FB-24, as will be seen in section 4, also a single seat version of the

FB-24 would have the same fuel capacity as the A-24 and with modern systems

becoming more autonomous the need for a second crew station may well diminish

before the FB-24 configuration is finalised.

Of the two configurations studied using AeroDYNAMIC V2.08 the NB2

configuration was demonstrated to have the best weight and hence lowest cost figures

therefore this configuration was further developed in the structural layout and systems

integration section below.

Further subsonic and supersonic analysis of range and endurance for both the

FB-24 and A-24 is given below in Appendices E at the end of this report pages

251 to 256, also it is worth noting that a single seat FB-24 would have the same

fuel capacity as the A-24 with the corresponding greater range and endurance,

and with enhanced single crew operation this should be considered for further

study.



148

The characteristics of the F-136-2 engine selected for this conceptual design study

were determined below from published data and installed thrust was calculated using

the methodology advised in reference 22.

There needs to be a much more comprehensive and detailed aerodynamic analysis

conducted on the NB2 configuration to enable more realistic assessment of the aircraft

configurations range and performance and ability to supercruise to be determined and

this is addressed as further work required in section 5 below.

For the F-136-2 engine with a published uninstalled maximum sea level thrust of

43,000-lbs would have a maximum installed thrust of 34,400-lbs.

The F-136 engines with an estimated uninstalled military power thrust of 34,400-lbs

at sea level would have a military power installed thrust of 27,520-lbs.

Using the uninstalled static sea level TSFC figures based on the experimental engine

i.e. TSFC at maximum thrust equal to 1.40[(lb/hr)/lbf] and at military power equal to

0.74[(lb/hr)/lbf]. Using the 20% factor for an installed engine these TSFC figures

become 1.68[(lb/hr)/lbf] for maximum thrust and 0.89[(lb/hr)/lbf)] on military

power.

Therefore for this project the installed sea level performance of the F-136 engine

selected for the Advanced Interdiction Aircraft FB-24 and A-24 derivative are as

follows:-

Maximum thrust = 34,400-lbs

TSFC at Maximum thrust = 1.68[(lb/hr)/lbf]

Military Power thrust = 27,520-lbs

TSFC at Military Power thrust = 0.89[(lb/hr)/lbf]

Taking the values for Maximum thrust / TSFC at maximum power and the Military

thrust / TSFC at military power calculated above, the fuel consumption for the

proposed mission can be determined for (1.1) the best cruse altitude (BCA) targeted

as 34,000ft and best cruse Mach (BCM) targeted as Mach 0.8, and (1.2) the high

speed Dash in weapon release and Dash out targeted at Mach =1.6 at an altitude of

45,000ft phases of the mission.

The S/L rate of climb rate envisaged for the AIA in the clean condition i.e. no external

stores and Maximum power is 50,000ft/min (254m/sec).

Using the equations for low bypass ratio turbofan performance modelling (Reference

22:- Introduction to Aeronautics: A Design Perspective: by Bradt. S: Stiles. R. J.:

Bertin. J. J: Whitford. R: Published by the American Institute of Aeronautics and

Astronautics, Inc: 1997) the TSFC over the mission portions of interest can be

determined as follows:-



149

TA = TSL ( / SL) (equation 1.1)

TA = TSL ( / SL) (1+0.7M) (equation 1.2)

Where: - TA = Thrust at altitude.

TSL = Thrust at sea level.

= Density at altitude.

SL = Density at sea level.

NB: - equation 1.1 is only valid for M < 0.9

ct = cSL (a / aSL) (equation 1.3)

Where; - ct = TSFC at altitude.

cSL = TSFC at sea level.

a = Speed of sound at altitude.

aSL = Speed of sound at sea level.

The installed sea level values being:-

Maximum thrust = 34,400lbs

TSFC at Maximum thrust = 2.51[(lb/hr)/lbf]

Military Power thrust = 27,520lbs

TSFC at Military Power thrust = 0.89[(lb/hr)/lbf]

Figures 97 and 98 below indicate some of the short comings of the F-35 when applied

to roles other than tactical interdiction and as will be seen below this conceptual

design has gone some way to address these issues, although its capabilities as a

bomber are an improvement over both the F-35 the F/A-22A and planned UCAV‟s

the latter being high subsonic aircraft it is not a fighter as typified by the F/A-22A in

figure 98 below, and lacks the agility for that role, which was not requested in the

request for proposals.



150

Figure 97: - Design limitations of the current F-35 family reflecting the need for

the AIA design divergence from the original baseline F-35C design.

Figure 98: - Comparison between F/A-22A and CTOL F-35 (published data),

contrast with figure TBD of AIA vs F/A-22A: JSF: Tornado: F-15E.



151

4.0 Structural layout and systems integration.

Figure 99: - F/A-22A internal structural layout illustrating structural density

source reference 17.

The objective of this section is to propose structural integration solutions for the

installation of the crew / AI system, propulsion system, offensive and defensive

weapons systems, undercarriage, and fuel systems within the FB-24 and A-24 OML

envelope, by designing a structure capable of accommodating these systems and

capable withstanding the aerodynamic and inertia loads to which these aircraft would

be subjected during the sizing mission. This section proposes the materials from

which the common FB-24 / A-24 airframe would be produced and highlights some

possible manufacturing routes. Each system is dealt with individually in sections 4.1

to 4.5, culminating in full structural integration modelling in section 4.6, which covers

layout, materials and weight.

Figure 99 above illustrates the structural density of advanced stealth combat aircraft

and the issues involved in systems integration. The level of definition in this figure is

much higher than could be achieved within this part – time MSc thesis and represents

an airframe post detail design which was beyond the scope of this work, however a

level of detail commensurate with the PWSC proposal submission phase was

produced here using notional skin thicknesses, and sub – structure thicknesses, and

the profile e.g. „I‟-section: „C‟- section are stated for each component or component

class. The complete airframe was solid modelled based on key datum face models

designed by the author which were intended for structural analysis using PATRAN /

NASTRAN FEA toolset however this analysis was not possible within the time scale

of this conceptual design study. Analysis of this structure was used to give a very

provisional indication of major structural component weight to support the proposal

submission. As stated above the NB-2 configuration was considered by the author to

have the best possibility of successfully meeting the design mission requirements and

was the only configuration carried forwards, although NB1 was used to draw some

comparisons with the NB-2 configuration.



152

4.1 Undercarriage systems integration for FB-24 and A-24.

Figure 100: - FB-24 and A-24 Common four tail airframe tip back angle = 18.70

for a wheel base of 8.086m and an aircraft C of G positions most fwd Frame

Station 9.19, (9.19m) and most aft Frame Station 10.11, (10.11m) along the a/c

axis.

Figure 101: - FB-24 and A-24 Common four tail airframe overturn angle = 75.30

for a wheel track of 3.325m, aircraft height of 4.71m and an aircraft C of G

positions most fwd Frame Station 9.19, (9.19m) and most aft Frame Station

10.11, (10.11m) along the a/c axis.

LG = 8.086m

CoG Most Fwd = FS 9.19 CoG Most Aft = FS 10.11

A/C height = 4.71m

18.70

Ground line

CoG Most Aft = FS 10.11

CoG Most Fwd = FS 9.19

75.30

W = 3.328m

53.50

420



153

Figure 102: - FB-24 and A-24 Common ruddervator airframe tip back angle =

18.70 for a wheel base of 8.086m and an aircraft C of G positions most fwd

Frame Station 9.19, (9.19m) and most aft Frame Station 10.11, (10.11m) along

the a/c axis.

Figure 103: - FB-24 and A-24 Common four tail airframe overturn angle =

75.250 for a wheel track of 3.325m, aircraft height of 3.79m and an aircraft C of

G positions most fwd Frame Station 9.19, (9.19m) and most aft Frame Station

10.11, (10.11m) along the a/c axis.

18.70



LG = 8.086m

Ground line

A/C height = 3.79m

W = 3.328m



71.90

53.50

420



154

The undercarriage used in this study and layed out above in figures 100 through 105

for NB1 and NB2 was a high Flotation / high Sink Rate type based on public released

data for the F-35C and measurements taken from X-35C drawings, and the main

characteristics of which were as follows and these were used to model the

undercarriage as shown above: -

Main undercarriage tiers sized at = 86.36 x 27.94 x 43.18cm:

Main undercarriage stroke = 33.17cm:

Nose undercarriage tire sized at = 59.69 x 19.05 x 25.40cm:

Nose undercarriage stroke = 52.07cm.

Analysis of the proposed undercarriage layout with a wheel base of 8.086m and a

wheel track of 3.328m demonstrated its suitability for both the NB1 four tail and the

NB2 ruddervator configurations, imparting a common tip back angle of 18.70 above

the minimum of 160 recommended in reference 21, and an overturn angle of 75.3

0 for

NB1 and 71.90 for NB2, both above the minimum angle of 62

0 recommended in

reference 21. The angle from the most fwd C of G position to the main undercarriage

oleo centre line was 53.50 for both NB1 and NB2, and the angle from the most aft C

of G position to the main undercarriage centre line was 420 for both NB1 and NB2.

The single nose wheel folds forwards into a nose wheel bay ahead of the oleo without

rotation about its axis, where as the main undercarriage also folds forwards into the

wing root blend but also rotates through 900 to stow into the wing blend wheel bay aft

of the wing leading edge spar as shown in figures 104 and 105 below, which is the

same mode of retraction used in the F-35 family of aircraft and thereby retains

functional commonality. In the event of a mechanical failure hanging up the

undercarriage the doors can be blown open with inert gas canisters and the

undercarriage can be lowered using a combination of inert gas and the airflow over

the aircraft underside into the locked position, thus enabling recovery of the aircraft in

such an emergency. The undercarriage bays were sized as: - nose wheel bay = 167cm

x 36cm x 54cm: and the main wheel bay = 288cm x 110cm x 52cm respectively.

The accommodation of the main undercarriage outboard of the fuselage in the wing

root gave a wide and stable wheel track enabling the possibility of deployment to

relatively austere forward airbases within NATO and allied nations, as well as good

handling in ice and poor weather which were primary mission requirements from

section 2. This layout did produce challenges however both aerodynamic in terms of

the extensive lower wing skin root to fuselage blend, and structural in terms of

incorporating the wheel bay and oleo within the root, and the design of the

undercarriage doors which needed to be both stiff and of complex curvature. There

could also be RCS issues arising from the need to eliminate any steps and gaps which

could arise from the articulated doors operation, although this would be addressed by

shaping the door with planform aligned chevrons and seals at the door periphery.



155

No

se

un

de

rca

rria

ge

retr

acts

forw

ard

th

rou

gh

94

0 in

to th

e

forw

ard

wh

eel

bay.

Main

u

nd

erc

arr

iag

e

retr

acts

fo

rward

thro

ug

h 9

20 in

to t

he w

ing

wh

eel b

ay

.

At

45

0

forw

ard

ro

tati

on

an

gle

th

e

main

wh

eel

itself

is r

ota

ted

th

rou

gh

90

0 a

rou

nd

th

e o

leo

to

lay f

lat

in t

he

wh

eel b

ay

.

Fig

ure

104:

- N

B2 U

nd

erc

arri

age

ret

ract

ion

an

d d

eplo

ym

ent

wh

ich

foll

ow

s th

e sa

me

met

hod

olo

gy a

s th

e F

-35 f

am

ily.



156

Figure 105: - Main undercarriage stowage in the wing root blend of the NB2

configuration, based on the F-35 family methodology.

4.2 Aircrew and AI integration for the FB-24 and A-24.

The as designed FB-24 NB-2 aircraft which was the final manned configuration

selected for this proposal submission enables the accommodation of two fully suited

flight crew representing the average 95th

percentile pilot in conditions required for

their completion of the desired mission. The original intention was to accommodate

these crew members in 1m high Martin Baker Mk-16 ejection seats, under a single

bow 2.6m canopy as envisaged for the stillborn F-22B trainer version of the F/A-22A,

however this would involve a complete redesign of the forward fuselage build module

and eliminate any commonality with the F-35 family of aircraft from which the FB-24

was derived, sharply increasing costs of manufacture, as well as increasing the drag

penalty of a canopy, and increasing structural weight. Therefore an alternative

solution was required, as a minimum change alternative and this was the direction

followed in this section for the manned airframe. In the region of the canopy there

were the conflicting requirements of drag reduction for supercruise flight favouring a

smooth cross-section distribution over the forward fuselage, and the need to provide

pilot visibility equal to that of the F-35 family for the Aircraft Commander (front seat

pilot) providing at least 150 of vision over the nose in level flight and ground

handling. Consequently there was a strong desire to retain the original F-35C layout

and substructure which gave the front seat pilot a published (reference 6) over nose

view angle greater than 160.

Wheels rotate through 900

to lay flat in the wheel bay, reducing storage envelope to a minimum in the x, z,

plane.



157

For the FB-24 an angle of 16.250 was obtained after some maturation and the design

eye location is shown in figure 106 below, in terms of FS (frame station) and WL

(water line) position on the aircraft axis.

Desig

n

ey

e

of

pilo

t air

cra

ft

co

mm

an

der

FS

367

0.3

WL

32

76.6

(Wate

r L

ine P

lan

e)

16.2

50

Vie

w a

ng

le

WL

21

35.7

(Wate

r L

ine P

lan

e)

Sta

tic g

rou

nd

lin

e.

Fig

ure

106:

- N

B2 D

esig

n e

ye

of

pil

ot

(air

cra

ft c

om

man

der

) w

ith

was

a 1

6.2

40over

nose

vie

w a

s in

th

e F

-35 f

am

ily.



158

This left the issue of the location of the second crew member the Mission Operation

System / UCAV Commander (MOSC) within the airframe, after dropping the tandem

canopy option shown in figure 107, and to provide this crew member with a means of

ingress / egress and emergency escape, as well as situational awareness.

The latter could be addressed by artificial vision and computer controlled imagery, the

MOSC would receive all information from the aircrafts onboard sensors, off – board

platforms e.g. AWAC, Recon UAV‟s / Global Hawk, and the synthetic real-time

environment of the A-24 UCAV‟s, and all of this information would be available in a

choice of single aspect or multi window formats on the 20” flat panel (God‟s Eye

View) touch sensitive display driven by a 400Gb 4.0GHz EMP hardened CPU

currently projected to be the size of a large notebook pc.

The former however proved more challenging, although after evaluation of the fuel

tank layout required to house the mission fuel the author determined that the MOSC

could be accommodated in the F-1 fuel tank location between the intake ducts, by

moving this fuel into the extended F-2 and F-3 tanks resulting from the fuselage

extension and removal of the internal gun provision (reversion back to the

missionized gun pod of the F-35C / F-35B variants), and the adoption of hose and

drogue refuelling as opposed to the USAF traditional boom method, covered in

section 4.4 of this thesis.

Figure 107:- The original crew accommodation layout under a large F-22B type

canopy, with height adjustable seating would have increased drag from the

canopy, and increased overall weight (generic aircrew and seats reference MB).

This location was supported by the fact that in the F-35 family the canopy uniquely

hinges forward just below the windshield as shown in figure 108, permitting the

fitting of a single chevron clam shell door behind the canopy for the second crew man

without risk of fouling the canopy hinge mechanism. The resulting structural

arrangement required to this is shown below in figure 109. The door would remain

sealed with the same type of pressurisation sealing as the canopy for the duration of

the mission, although pyrotechnic locking bolts would provide quick door jettison into

the airflow down stream over the fuselage permitting safe ejection of the MOSC in an

emergency.

2.6m

THE ORIGINAL TWO CREW STATIONS



159

Figure 108: - The F-35 families‟ unique forward pivot canopy hinge mechanism

which is shown here on the F-35A airframe mock-up model (reference 28).

Figure 109: - Crew accommodation integration in the FB-24 showing extent of

frame cut out to house second crew station in space envelope of F-1 fuel tank,

modelled by the author in CATIA V5. (Note substructure is complete consisting

of frames and longerons thickness modelled is 50mm (flange width) where as

actual frame pocket thickness will be 4mm in carbon PMR-15.)

Orange frame segments are part of clam shell door.

Pilot (aircraft commander)

MOSC (UCAV commander)

16.250

Door hinge frame

3.5m



160

The seat pitch in the FB-24 was set at 220 as in the current F-35C aircraft design, as

there was no requirement for the crew to endure loads over 7g throughout the duration

of the mission. The seat dimensions were 1.20m high by 0.50m by 0.54m

commensurate with the published dimensions for the Martin Baker Mk-16 seat

designed for the F-35 family. The 0.61m3 volume requirements for the crew systems

i.e. OBOGS, and OBIGS could also be contained within the as designed structure

shown in figure 109 above outboard of the MOSC crew station, within the DIS IML

boundary, although definitive positioning would undertaken as further work under

detailed systems integration. The design of the clam shell MOSC door was also left to

follow on work due to the time constraints of this study, but the author envisages a

cover which maintains OML continuity with the airframe. This concluded the FB-24

air crew integration study.

The A-24 AI (artificial intelligence system) integration was a simpler task, as this

system was envisaged as a complete self contained module with its own ECS as in the

X-45C and X-47B UCAV‟s requiring only a plug and play power interface, and fibre

optic sensor data link. The module installation did not adversely effect the provision

of the F-1 fuel tank and as such the A-24 fuel capacity was greater than that of the

FB-24 as will be covered below. Figure 110 illustrates the AI module installation as

modelled by the author in CATIA V5.

Figure 110: - AI installation integration in the A-24 showing extent of frame cut

out to house system space envelope without affecting the F-1 fuel tank, modelled

by the author in CATIA V5. (Note substructure is incomplete consisting of

carbon PMR-15 frames only full structure is shown in section 4.6).

However the height of the module i.e. 1.25m did break the fuselage OML if the

canopy envelope was to be replaced with a fairing to continue the curvature of the

surrounding skin.

AI Module Optical

tracking system

AI Module with ECS incorporated

F-35 Standard radar

mounting frame

1.74m

Carbon PMR-15 all frames



161

Therefore with the module in its current form the canopy shape will need to be

retained with its optical properties. However the drag penalty will be off - set by the

fuel capacity of the F-1 tank, as discussed in section 4.4. The canopy could therefore

serve as the primary access and servicing aperture, and because it is already designed

for low RCS reflections and emissions there will be no adverse affects on the aircraft

signature from the canopies retention. This concludes the section on the aircrew and

AI system integration, in which the ability to accommodate two aircrew within the

substructure and skin IML was established, although further work is required to

define the MOSC access door. The ability to accommodate the AI system in the

systems current preliminary design form was also established and no adverse impact

on fuel capacity was determined.

4.3 Propulsion systems integration for the FB-24 and A-24.

The F-135 Propulsion system: - The F-135 propulsion system has been further

developed over that of the JSF119 shown in figure 111 and installed in the X-35

Concept Demonstrator Aircraft (CDA) currently the F-135 figure 112, has been rated

at 43,000-lb uninstalled thrust with afterburner, and has bypass ratio in the order of

0.30 and with 67% of maximum thrust available in military power the dry thrust was

calculated to be 28,810-lb. The engine integrates the proven F-119-100 core and has a

three – stage fan with hollow first stage blades, and composite fan inlet guide – vanes.

Behind this fan section is a six – stage high pressure bladed blisk compressor, a single

– stage high pressure turbine, and a two – stage low – pressure turbine, which has a

new low – pressure spool replacing the single - stage of the F-119 engine. In addition

the propulsion system features advanced prognostic and on – condition management

systems that provide maintenance awareness, automatic logistic support, and

automatic field data and test systems. This engine also has 40% fewer parts combined

with 50 % lower infrastructure support requirements compared with current in service

engines, according to the manufacturer.

The variant considered for the FB-24 and A-24 would be the F-135-400 CV variant

engine as this engine is proposed to have an installed thrust of 43,000-lb with

afterburner.

Figure 111:- Pratt & Whitney JSF119 low by – pass ratio turbofan cut – away,

showing the very high degree of F-119 commonality source P&W press office.



162

Figure 112:- The Pratt & Whitney F-135-100 JSF CTOL engine under final

assembly at Pratt & Whitney‟s plant, and like the F-119 it was designed to be

maintained with six common hand tools and the elimination where possible of

wire lock fasteners. (Source: - Reference 6.)

The F-120 / F136 Propulsion system: - The General Electric F-120 engine program

was selected by the JSF Joint Project Office as the baseline from which to develop an

Alternative Engine Program (AEP) to the Pratt & Whitney F-135.

Figure 113:- General Electric F-120 Variable Cycle Engine Schematic illustrating

the engines two modes of operation (Source reference 17).



163

Figure 114:- Longitudinal Section showing main component layout of the General

Electric F-120 Variable Cycle Engine (Source reference 26).

This selection was based on the F-120 engines core thrust capability to meet the JSF‟s

multi - service aircraft requirements, including STOVL capability without scaling up,

and for its growth potential. Originally designated the YF-120FX for the JSF, the

engine is based on the GE F-120 developed and successfully demonstrated in flight in

the early 1990‟s in the USAF Advanced Tactical Fighter engine competition.

Development of the F-120 engine shown above in figures 113 and 114 has continued

since the ATFE competition under the Advanced Research Projects Agency‟s

advanced short take-off and vertical landing program which has seen the engine run

for 21 hours, and since 1995 funding has come from the JSF Program Office, for: -

Phase 1 engine definition (1995-1997): Phase 2 Critical Design Review (1997-2001):

Phase 3 Detail design (2002-2005) and 12,000+ ground test hours. Currently an F-136

derivative F-120 engine planned to fly in 2009 with first production engine delivery

anticipated 2011 and the engine was formally designated the F-136 in 2005, however

with the on going drain on the US defence budgets this was reassessed and after the

Quadrennial Defence Review of February 2006 the possibility exists that this

alternative engine may be delayed or cancelled. This thesis assumes that funds will be

found for this engine within the development time frame of the FB-24 and A-24

aircraft i.e. 2015 through 2020.

The technical characteristics of the F-136-2 are as follows:-

Front fan (Rolls-Royce): Long wide - chord, titanium, three–stage blisk: stage one

has hollow core blades: while stages two and three have solid blades. (Currently two

builds have been tested, which has verified the fan flow, as well as the efficiency of

linear friction welding for blade attachment).

High-Pressure compressor (General Electric GE): Five – stage, all – blisk system

consisting of three rotors in stage one and stage two, and stages three to five being

inertia – welded together.

No

zzle

rep

laced

by

LO

AN

on

FB

-24

an

d

No

rth

rop

n

ozzle

on

A-2

4.



164

The compressor features forward swept airfoils, and robust blade tips, bowed / swept

stators designed from 3-D aerodynamic codes, and high stage loading to support

40,000-lb thrust class operation.

Combustor (Rolls-Royce): Single annular simplified design, fabricated from

Lamilloy cooling material, and based on the IHPTET (Integrated High Performance

Turbine Engine Technology) program research into higher turbine operating

temperatures.

Turbine (GE and Rolls-Royce): Single-stage high – pressure turbine (HPT) and a

three stage low-pressure turbine (LTP), the HPT and stage one LPT forming a

coupled, vaneless counterrotating system (with single crystal HPT blades). This

turbine system has been successfully rig - tested at the time of writing.

Augmentor (GE): Radial, non-stage, variable flow control, based on GE YF-120:

F414-GE-100: F110-129 and F110-132 engines.

Technical information from the following General Electric Aircraft Engine web page:

- http://www.geae.com/engines/military/f136/background.html accessed on the 1st

February 2005.

The original YF-120 design employed a fan and airflow size 12% larger than that of

the XF-120 experimental ATF engine. This size increase was in response to the

requirement for more cooling air for the metal exhaust system and the increased thrust

requirements. Overall pressure ratio was 22. Durability was addressed using thermal

barrier coatings (TBC) and tailored cooling air distribution.

As a result of the engines short and compact hot section vaneless high – pressure / low

– pressure turbine concept, the hot section cooled surface area was 30% less than in

the General Electric F-110 engine. Variable cycle engine (VCE) features were

simplified from the XF-120 design, and maintainability requirements were rigidly

enforced during the design process. At time of writing this thesis the F-136 has been

rated at 43,000-lb uninstalled thrust with afterburner, and has bypass ratio in the order

of 0.32. For the original YF-120 engine dry thrust was 28,000-lb using VCE

technology as shown in figure 113 were the engine can be operated as either a turbofan

or a turbojet when the maximum thrust was 35,000-lb or 80% of maximum thrust was

available in the dry condition. Translating this into the current figures for the F-136 if

the same VCE technology was employed then for a 43,000-lb maximum thrust a dry

thrust value of 34,400-lb could be achieved.

For both the F-135 and F-136 engines all components, harnesses, and plumbing were

located on the bottom of the engine for easy access, and all line replaceable units

(LRU‟s) were located one layer deep (i.e. units were not located on top of each other),

these features and a reduction of the tools required to maintain the engines to just six

standard tools increases the supportability of the engines in service.

The basic engine dimensions provided by GE and PW press offices that were

used in this thesis are given below:-

For both engines the fan face diameter was = 1.10m (3ft 7in):

http://www.geae.com/engines/military/f136/background.html



165

For both engines the overall length including nozzle was = 6.06m (19ft 10in):

The nozzle length for both installations was = 0.72m (2ft 4in):

The nozzle exit face diameter was = 0.79m (2ft 7in):

The nozzle base diameter was = 1.27m (4ft 2in)

The overall engine weight including nozzle and equipment was = 3,402kg

(7,500lb) based on the F-119 engine weight figures.

Figure 115: - Proposed FB-24 and A-24 Propulsion system integration in which

the standard bifurcated intake duct of the F-35C is integrated with the General

Electric YF-120 derived Variable Cycle Engine (F-136-2) and GE Axisymmetric

Vectoring Nozzle (AVEN). This installation was modelled by the author in

CATIA V5, and the structural integration was modelled in section 4.6.

The FB-24 and A-24 utilise the F-35 family bifurcated composite intake ducts and the

diverterless supersonic intake DSI (see appendices C) as shown above in figure 115.

The fuselage extension for wave drag reduction had the additional benefit of enabling

a longer flow recovery stage to be introduced between the intake convergence point

and the engine face reducing the risk of turbulent airflow and shockwaves causing

flow separation and pressure loss reaching the fan face, thereby permitting better

pressure recovery at Mach 1.6. That is the flow recovery stage of 3.7m for the FB-24

and A-24 compared with 1.7m for the F-35 family.

6.06m

6.45m

2.93m

YF-120 derived F-136-2

Variable Cycle Engine

AVEN Nozzle

Bifurcated intake duct

Aircraft Centre Line and Engine Thrust Line.

Fixed trunnion

Free sliding trunnion

Fwd link



166

An example of the bifurcated made from CFC without the use of mechanical fasteners

is shown below in figure 116. The bifurcated intake results in a 100% line of sight

blockage of the engine face and thus eliminates all return energy leaks substantially

reducing the FB-24 and A-24 RCS which was in common with the F-35 family of

aircraft.

Figure 116:- F-35 / FB-24 / A-24 common intake duct design with the supporting

stiffeners being co-bonded and z pinned to the duct wall thus obviating the need

for mechanical fasteners thereby reducing weight and potential RCS reflection

sources. (Source: - Lockheed Martin JSF media relations).

Engine installation and removal would be via the titanium engine bay doors located

on the underside of the rear fuselage, as per the current F-35 family as shown in figure

119 below. The engine locates onto an integral rail machined into the top of the rear

fuselage frames which with side mounting bolts transfer the thrust loading to the

airframe, and the engine will be screened by refractory coated Niobium alloy or

carbon BMI nacelle skins from the rear fuselage substructure, as shown in figure 117,

details of this structural arrangement for the F-35 are ITAR restricted and therefore

cannot be elaborated on within this thesis. The engine bay will have a full fire

suppression system. For transportation to and from the aircraft the engine the common

A/M32M-34 trailer used for both F-35 and F/A-22A aircraft shown in figure 118,

which is 4.26m (14ft) in length and 1.8m (6ft) wide, and 1.5m (5ft) high with a

maximum payload of 3,402kg (7,500lb).



167

Figure 117:- CGI modelling of the installation of the F-136 engine into the F-35C

from the common A/M32M-34 engine transport trailer which can be raised and

lowered as shown. The identical system would be used for the FB-24 and A-24

aircraft. (Source Lockheed Martin F-35 Project Media relations).

Figure 118:- Proposed engine mounting for FB-24 / A-24 based on F-35 family

(Source reference 6).

Engine attach rail.

Engine free - sliding trunnion mount.

Engine fixed trunnion mount.



168

Figure 119:- The Boeing developed A/M32M-34 trailer which greatly eases the

installation and removal of the F-136 / F-119 / and F-135 engine by permitting

extremely precise vertical and lateral movement of the engine. (Source: -

Reference 14)

Engine mounting within the airframe was achieved with the addition of a minimal

amount of interconnecting structures, but the engine mount had to be able to prevent

deflections in the main aircraft fuselage structure from inducing loads into the engine

itself, and also had to allow for the thermal expansion of the engine both along its axis

and around its circumference. The basis for the method used for the FB-24 and A-24

and shown above in figures 115 and 118 was based on the Lockheed F-104 Starfighter

fighter aircraft mounting system described in reference 28: - Airframe Structural

Design (Practical design information and data on aircraft structures): by Niu.

Michael. C. Y.: Published by Hong Kong Conmilit Press Ltd 1997, because the F-104

was a single engine fighter in which the engine was mounted in the rear of the

fuselage, and was integrated with bifurcated ducts.

The major portion of the vertical loads shown below in figure 120 would be carried

by the trunnions located close to the engines centre of gravity. Side (or lateral) loads

would be taken out on the fixed starboard trunnion only, with the port trunnion being

free to move laterally to allow for any radial thermal expansion of the engine. The

forward mount would be a universal joint capable of carrying vertical loads only.

Because the would be located near the C of G the major forces experienced by the

forward support link mount would come from the gyroscopic couple crated by the

angular velocity in yaw and the inertia moment caused by angular acceleration in

pitch which would be relatively small.



169

Unlike piston engines the moving components of the gas turbine engine have a simple

rotary motion and the combustion process is continuous as apposed to intermittent as

in the case of a piston engine thus the unbalanced forces which could result in

mechanical vibrations would be small in size and few in number, therefore

mechanical vibration should not be a major issue.

Figure 120: - The F-136-2 engine / fuselage mounting arrangement based on that

of the Lockheed F-104 Starfighter, illustrating the loads carried by the engine

mounts, as explained above.

Acoustic vibration on the other hand resulting from the jet plume could have had a

major impact on the down stream empennage booms which had to have inboard faces

angled outboard by 150 to avoid the nozzle acoustic wave cone, because prolonged

exposure would result in acoustic fatigue. This angling outboard of the inboard boom

faces was used on the F-35 family aircraft to reduce potential for acoustic fatigue, for

the short stock booms to which the horizontal are mounted, by comparison the FB-24

/ A-24 booms were nearly twice as long and tapered so the risks were perceived to be

grater for this aircraft.

Although the AVEN nozzle was a thrust vector nozzle the potential for acoustic

damage resulting from its use off the aircraft centre line was considered slight because

of the very short dwell times involved (fractions of a second on average) to invoke a

manoeuvre, therefore the current layout was deemed adequate to withstand any

impingement.

This concludes this thesis coverage of the engine system integration, and the

integration into the fuselage substructure is shown in section 4.6 below (to the extent

to which this can be reported).

Free - sliding trunnion mount for engine radial

expansion.

Fixed trunnion mount.

Link – mount for engine

axial expansion.

Vertical

Forward Lateral

F-136-2 Engine



170

4.4 Weapons systems integration for the FB-24 and A-24.

This section covers the accommodation of the primary and probable future weapons

within the FB-24 and A-24 airframes and the resultant sizing of the internal weapons

bays, as well as the installation of the avionic systems required for target acquisition

(within ITAR restrictions) and the proposed defensive systems for these aircraft.

The primary weapons for these aircraft called up in the requirements document were

as follows:-

JDAM GBU-31PIP

Release weight: 959.3kg (2,115lb) Length: 3.77m (12.38ft) Max span 0.6m (2.1ft) Cost $20,000

ALOSNW (Projected weapon for FSAV) Release weight: 1133.9kg (2,500lb)

Length: 4.26m (14ft) Max span: 0.45m (1.5ft)

Warhead: Evolved W69 (200kt‟s)

Air-to-Air Weapons

ASRAAM Missile Launch weight: 87kg (192 lb) Length: 2.8m (9.51 ft) Max span: 0.48m (1.6 ft) Body diameter: 0.16m (0.54 ft) Launcher rail weight: 22.8kg (50 lb) Launcher rail length: 2.8m (9.5 ft)

The M61A1 20mm cannon requested in the requirements documentation was to be

fitted as an optional weapon in the form of a missionized pod external to the airframe

structure as per the F-35C, and the F-35B aircraft, in order to increase internal fuel

provision and simplify the structural design of the forward fuselage reducing

structural weight. This was considered as acceptable by the customer as the internal

provision for this weapon was solely based on legacy air superiority aircraft and had

no specific relevance to either the FB-24 or A-24‟s primary mission.

The ALOSNW was comparable in size weight and shape to the JASSM, figure 121 a

conventional weapon and this latter weapon was used to size the weapons bay volume

for the former as a public domain system which can be openly discussed within this

proposal.



171

Figure 121: - JASSM mounted on a storage / transportation fixture as a

conventional low observable stand – off attack munition with a range of 200nm

but the same OML and dimensions as the ALOSNW this was used sizing and as

a future store for the FB-24 / A-24 weapons system. (Source Lockheed Martin

ADP).

AGM-158A / B JASSM is a 2,250lb turbojet powered cruise missile with a range of

more than 200nm, and a 1,000lb warhead multi-purpose warhead effective against

hardened targets, and a light weight composite airframe. This weapon uses GPS to

navigate to the vicinity of the target, where upon an IR sensor in the nose images the

target and the weapons computer system compares it to a visual template loaded

before the mission and based on reconnaissance images. This weapons small size,

planform aligned shape, and selective use of RAM, combine to make this a very

effective low RCS weapon which is reasonably cheap at $500,000 each. The JASSM

would be carried internally in common with all other FB-24 / A-24 expendable

weapons because as stated above neither of these aircraft have external weapons

provision as a result of the need to reduce airframe drag and maintain a low radar

cross section, both of these requirements could not be met from the external carriage

of stores. (Reference 29: - Ultimate Fighter Lockheed Martin F-35 Joint Strike

Fighter: Sweetman. Bill: Zenith Press USA: 2004).



172

The Boeing GBU-39/B Small Diameter Bomb low cost precision strike weapon

shown in figure 122 below, was also required to be carried by the FB-24 / A-24

aircraft in order to carry a greater number of weapons internally than two JDAMS,

and sizing information for this was obtained from Boeing (Reference 30:- Marguerite

Ozburn Global Strike Systems, Boeing IDS Business Support, Communications and

Community Affairs, P.O. Box 516 St Louis MO 63166 e-mail), and given below: -

Small Diameter Bomb GBU-39/B

Length: 1.8m (5ft.9”)

Width 19cm (7.5”)

Weight 130kg (265lbs)

Warhead Penetration 3ft of steel reinforced concrete

Range >11.12km (60nm) with wings

BRU-61/A “smart” pneumatic carriage

Payload capacity four GBU-39/B Small Diameter Bombs:

Weight empty = 145kg (320lb): Max weight = 664kg (1460lb):

Dimensions: - length = 3.6m (143”): width = 40.6cm (16”) height = 40.6cm (16”):

Figure 122: - Installation of the BRU-61/A “smart” pneumatic carriage with four

GBU-39/B Small Diameter Bombs into a weapons bay possibly a B-1 Lancer.

(Source Boeing IDS Global Strike Systems)



173

Figure 123: - Proposed ASRAAM launch method from the FB-24 / A-24 weapons

bay by retractable launch rail concept was based on the side weapons bay AIM -

9M/X ejector system of the F/A-22A using swing arms to deploy a stub rail

mounted inboard of the outer weapons bay hinge line (authors CATIA V5 model

based on references 14 and 29 no components sized except ASRAAM).

The ASRAAM infra red seeking short range missile would be carried for self defence

and was not considered as an offensive system. The weapons location within the

weapons bay rather than the external carriage as is the case on Tornado resulted in a

limited field of regard for the seeker head, being bound on one side by the weapons

bay door and above by the FB-24‟s lower fuselage, this being a different case from

the F/A-22A where the AIM-9 infer red seeker is swung out from fuselage side

weapons bays in clear air with a much better field of regard for the seeker head.

Therefore in the case of both the FB-24 and A-24 the ASRAAM‟s would be launched

in “Lock-on After Launch” (LOAL) mode which was the only option from the

weapons bay location. There would be issues with this approach because in short

range combat the target aspect could change dramatically and unpredictably, and by

the time the missile clears the airframe the target may not be where it was at firing

initiation, and something else may have assumed that location e.g. the FB-24‟s

wingman, so time could be lost in trying to re-acquire the target. These issues would

be addressed to some degree by off board tracking and data feed and further research

would be required to build confidence in this deployment.

As explained above in figure 123 the ASRAAM would be swung out of the weapons

bay on a stub launch rail using actuated swing arms mounted on the weapons bay

internal longeron / outboard side wall, this was in preference to using a door mounted

rail as the door would have become too heavy if stiffened to carry the rail, which

would have increased the size of the door actuators and surrounding structure. With

this system the ASRAAM exhaust plume is clear of the weapons bay and little

additional stiffening of the bay structure would be required.

Door

Bay longeron

ASRAAM

Swing arms

Actuator

Bay door hinges

Launch rail



174

Figure 124:- The port side weapons bay of the F-35A common to the F-35 family

with one 2,000lb JDAM installed (Source reference 29).

Figure 125:- Illustrates the results of the first test in the FB-24 / A-24 redesigned

F-35C which shows that the proposed lager weapons bay of the FB-24 could not

be accommodated in the positions indicated for the F-35C weapons bay doors on

the Scalecraft model which appear to be incorrect. The conclusion was to move

the new bays aft which was closer to the aircraft C of G and beneficial.

Initial 4.5m long 0.63m wide 0.5m high weapons bay in F-35 model weapons bay door location impinges on intake duct, and therefore was rejected.

STOVL model 3.5m long 0.63m wide

0.5m high weapons bay is a snug fit.



175

Figure 126:- The final weapons bay integration aft and clear of the bifurcated

intake duct with the doors starting at Frame Station FS-5905.9 and ending at

Frame Station FS-1155.6

After sizing studies using the data in table 2 page 34 and data from reference 15, as

well as fit checks in CATIA V5 (figure 125) and with 1/48 scale kit models of the F-

35 the weapons bay was sized at: - length = 4.55m: width = 0.64m: height = 0.51m

giving a volume of 1.48m3 for each bay which was sufficient to accommodate one

ASRAAM plus one JDAM GBU-31 PIP, or one ASRAAM plus one JASSM, or one

ASRAAM plus one ALOSNW per bay. Also up to six GBU-39B Small diameter

bombs per bay could be accommodated without the BRU-61 rack or four per bay with

the rack but without the ASRAAM missile or launch rail. These weapons bays met

the operational requirements captured in section 2 of this thesis, and their integration

into the extended fuselage is shown above in figure 126, achieved with analysis of

public release data which sites the weapons bay doors too far forward on the baseline

F-35C in relation to the bifurcated duct leading to a clash which was resolved by

moving the bays to start at Frame Station FS-5905.9, and close out at FS-1155.6.

In addition to the primary weapons currently in service or approaching service entry

the author was asked to consider two future weapons which will enter service by the

entry date of the FB-24 / A-24 airframes, which were the LOAAS battlefield loiter

attack weapon figure 127 and the SACM small cruise missile figure 128. From the

data provided both weapons can easily be accommodated within the weapons bay and

expelled using the same cold gas ejector system as would be used for the FB-24 / A-

24 primary air to ground weapons.

In all cases of air to ground weapons release the author proposed that a drop down and

retractable flow disruptor plate was installed at the front of each weapons bay to aid

clean separation of the store from the aircrafts flow field, so that the stores fall away

and are not captured in the weapons bay.

10.91m

1.78m

0.64m

4.55m Volume 1,48m

3



176

Figure 127:- One of the future weapon considered was the LOAAS battlefield

loiter GPS guided weapon: length 0.91m (3ft): weight 45kg (100lb): This would

be ejector launched from the weapons bay, and is powered by a 13kg (30lb)

thrust turbojet engine, LOAAS has a straight line range of 100nm and a

15minute duration at 750ft and 200knts covering a foot print of 25nm2. (Source

Lockheed Martin ADP Skunk Works)

Figure 128:- The other future weapon considered was the SMACM small attack

cruise missile which can be used both as a reconnaissance or attack platform:

length 1.7m (5.8ft): weight 64kg (142lb): This would also be ejector launched

from the weapons bay, and is powered by a J45G turbojet imparting a range of

200nm at BCA/BCM. (Source Lockheed Martin ADP Skunk Works)

This concludes the weapons bay integration study for this thesis, and further work

required is highlighted in section 5.



177

4.5 Structural layout of the FB-24 and A-24.

The objective of this section was to present the airframe structural layout maturation

for the NB2 configuration and the reasoning behind the design selections for the final

PWSC FB-24 / A-24 conceptual design submission.

This section covers structural layout description, structure to structure joint definition,

major component attachment selection, build joints, selection, as well as the material

selection to be used in the manufacture of the primary structural members. The

development of each major structural component was described in turn starting with

the wings, followed by the forward fuselage, centre fuselage aft fuselage and

ruddervators and culminating in the full airframe structural model.

In order to establish an inner mould line IML the author used notional constant skin

thicknesses based on the public domain values for legacy aircraft namely the F/A-22A

and F/A-18E Super Hornet (Reference 31: - Lessons Learned from the F/A-22 and

F/A-18E Development Programs RAND_MG276: by Younossi. Obaid: Stem. David.

E. et al: Prepared by for the USAF by RAND Project Airforce), and from the work of

Alan Baker et al in Composite Materials for Aircraft Structures (Reference 32: -

Composite Materials for Aircraft Structures second edition: by Alan Baker et al:

Published by the American Institute of Aeronautics and Astronautics 2004) nominal

skin thickness values of 5.08mm for the fuselage and 20mm wing root tapering to

4mm at the wing tip were used for all structural modelling. Due to the time constraints

involved in this conceptual design all frames, spars, and ribs were modelled with

thicknesses representative their respective flange widths not their web thicknesses

therefore all spars are modelled as 76.2mm, ribs are modelled as 63.4mm, frames as

50mm, longerons as 30mm, and keels as 50mm thickness. Thus the structure looks

over designed however the intention is to impart layout and not detail design to the

reader. Also due to the time constrains the structural joints between members were not

modelled but their type is indicated as either “Bath-tub” or “Tag on stiffener” were

possible no Cleats were to be used to reduce parts count.

The airframe structure was conventional for a modern interdiction aircraft employing

high-speed machined aluminium, near net forged titanium, and HIP processed

titanium, carbon BMI skins throughout, unlike the F-35 in which most of the wetted

area was carbon epoxy, and some aluminium honeycomb core stricture. The only

fabrications envisaged are the engine bay doors and the land based arrestor cover both

would be made from SPF/DB titanium. The structure was to be mechanically fastened

using the same build philosophy as the F-35C family thereby reducing cost of

ownership by ease of maintenance in service. Where possible handed parts have been

avoided and common non-handed parts developed for ease interchangeability of

components and structure. This extensive use of metallic substructure is due to two

key factors covered in references 30 and 31, namely the relatively high cost of

composite substructure compared to high-speed machined aluminium, and the formers

limited tolerance of ballistic damage. The major component build joint philosophy

was soft mate where skins, longerons and keels land on a single interface frame this

reduces structural weight compared with the more traditional hard frame to frame

mate joints employed on some legacy aircraft.



178

4.5.1:- Wing structural layout and undercarriage integration.

Figure 129:- F-35 original single component aluminium and titanium wing

structure which was mated at a water line joint with the centre fuse frames,

resulting in a heavy and complex joint and this has been changed on all SDD

aircraft after AA-1. (Source Lockheed Martin Public affairs office)

Figure 130:- F-35 original single component carbon epoxy thermoset wing top

skin produced by fibre placement, the wing skin also incurred weight growth

with maturation although the concept was good from a signature reduction stand

point. This has also been revised in all SDD aircraft after AA-1. (Source

Lockheed Martin Public affairs office)



179

Figure 131:- F/A-22A wing torsion box illustrating a two component wing

structural layout with spars connected to single piece frames, and is composed of

titanium and carbon composite spars, a lighter solution than that originally

selected for F-35. (Source Lockheed Martin Public affairs office)

Figure 132:- F/A-22A Carbon BMI wing skins which have proven resilient in

service to the supercruise environment, and these are lighter by area than the

original single component wing after the latter design had matured. (Source

Lockheed Martin Public affaires office)



180

Figure 133:- Phase 1 initial FB-24 / A-24 wing structural layout overview

detailed in the text below and in figure 136.

The design choices available for wing major structural component of the FB-24 / A-

24 common airframe as shown the figures 129 through 132 above and were a

continuous wing torsion box with a centre fuselage waterline frame break shown in

figures 129 and 130 favoured for the F-35 - 230-5 PWSC design and built as F-35

SDD aircraft AA-1, or two separate wing components (port and starboard) attached to

continuous centre fuselage without any frame breaks shown in figures 131 and 132

which has been successfully used on the F/A-22A now in service with the USAF.

Three factors drove the author to select the latter wing design approach, which were: -

(1) Complexity of the split frame joint: (2) Weight implications on the fuselage

frames: (3) Weight implications for the continuous wing substructure. Unlike the

BAE Systems / Boeing Harrier where the wing sits on top of the frames attached by

six pick-up fixtures, the F-35 wing mate joint was at 40% of the fuselage depth, and

although distributed over a lager number of frames would still require substantial

thickening in these members to transfer the wing bending and shear loads from the

spars and the fuselage bending loads from the longerons and keel. Additionally apart

from the two Concept Demonstration Aircraft X-35A/B and X35C this depth of joint

had no in service reference on legacy aircraft and was considered high risk by the

author.

Therefore the more conventional design philosophy shown in figures 133 and 134 was

selected for the common FB-24 / A-24 wing, with a metallic substructure of titanium

peripheral members and an advanced aluminium alloy core substructure. The skins

were carbon BMI (Bismaleimide matrix) high temperature thermoset with Tg values

ranging from 1750C to 235

0C although normal operational temperatures for long

periods in the hot wet condition would usually be limited to 1500C (reference 31).

This material was also a proven material for the F/A-22A wing skins which operates

in the same environment envisaged for the FB-24 / A-24 airframe. The wing structure

is detailed below in figure 134.

Ti boundary structural members

Ti root rib

Al core sub -structure

Main undercarriage bays

Carbon BMI skin 5.52m

9.39m



181

Fig

ure

134:-

Det

ail

ed d

escr

ipti

on

of

Ph

ase

1 w

ing s

tru

ctu

ral

layou

t.

Tit

an

ium

L

ead

ing

ed

ge

sp

ar

ballis

tic

d

am

ag

e

resis

tan

t w

ith

exce

llen

t fa

tig

ue r

esis

tan

ce

Tit

an

ium

Lead

ing

ed

ge

flap

att

ach

men

t ri

b b

ath

tub

jo

inte

d t

o s

par

Tit

an

ium

Lead

ing

ed

ge f

lap

att

ach

men

t ri

b b

ath

tu

b jo

inte

d t

o s

par

Tit

an

ium

Aft

sp

ar

wit

h

bath

tu

b

join

ts

to

all

att

ach

ed

str

uctu

re

Tit

an

ium

Tip

rib

Tit

an

ium

Fla

pero

n

att

ach

men

t ri

b

Tit

an

ium

Fla

pero

n

att

ach

men

t ri

b

Tit

an

ium

Ro

ot

rib

bath

tu

b jo

ints

to

all a

tta

ch

ed

str

uctu

re

Alu

min

ium

co

nti

nu

ou

s m

ain

sp

ar

bath

tu

b jo

ints

to

rib

s

2 p

art

Alu

min

ium

sp

ars

bath

tu

b jo

ints

to

rib

s

2 p

art

Alu

min

ium

rib

1

bath

tu

b jo

ints

to

sp

ars

3 p

art

Alu

min

ium

rib

2

bath

tu

b jo

ints

to

sp

ars

Alu

min

ium

cap

rib

3 b

ath

tu

b jo

ints

to

sp

ars

Wh

eel

bay

A

ctu

ato

r b

ay



182

Figure 135:- Phase 2 maturation FB-24 / A-24 wing structural layout overview

detailed in the text below and in figure 138.

The initial Phase 1 layout shown in figures 133 and 134 formed the basis for all

subsequent wing maturation studies. The author had concerns that the Phase 1 spar

pitch was too wide leading to large skin panels which were susceptible to buckling,

therefore three additional sub spars were added in Phase 2 maturation to break up the

Phase 1 skin panels into thinner panels which would be more resistant to buckling

(reference 31), as shown above in figure 135 and detailed below in figure 136. The

number of ribs was not increased as their function was to transfer the flaperon loads

and leading edge flap loads into the spars and the Phase 1 layout was considered

capable of achieving this.

During Phase 2 the wing attachment philosophy was selected based on reference 28,

chapter 8. From the types of wing root joint described in this work the most suitable,

and that with which the author has had first hand experience of was the shear lug type,

which were easily assembled, least costly to manufacture compared to splice plates,

and generally more economic for military aircraft with thin airfoil sections such as the

FB-24 / A-24 compared to transports. This system was also used on the F/A-22A so

the author considered that the weight penalty alluded to in reference 28 should not be

sever enough to exclude this type from consideration for the FB-24 / A-24 aircraft.

In the design presented in this thesis both top and bottom lugs are in the same plane,

and both take axial loads, with the vertical load being shared between them (although

this is difficult to predict, and the small moment arm produces high lug axial loads).

The major advantages were a strong fitting at relatively low machining cost.

New Ti structural members

New Al spars

Moment fittings

Fwd shear fitting

Aft shear fitting



183

All of these fittings were to be titanium so a logical solution was to make all spars

below Rib 1 titanium and split the Root rib into sections which would have bath-tub

joint into the spars as shown in figure 136 below.

Fig

ure

136:-

Det

ail

ed d

escr

ipti

on

of

Ph

ase

2 w

ing s

tru

ctu

ral

layou

t.

Ti

Lead

ing

ed

ge

sp

ar

wit

h

inte

gra

l sh

ea

r att

ach

men

t.

Al

/ T

i R

ib

1

Fla

pero

n

att

ach

men

t b

ac

ku

p r

ib a

nd

als

o

the f

uel

tan

k b

ou

nd

ary

wit

h t

he

wh

eel b

ay

bro

ken

by r

ear

sp

ar

Al

Rib

2 F

lap

ero

n c

en

tra

l att

ach

men

t

rib

sp

lit

by

sp

ars

4, 5,

an

d 6

Ti

Stu

b

sp

ar

1

wit

h

inte

gra

l

sh

ear

/ m

om

en

t att

ach

me

nt.

Al

/ T

i S

tub

sp

ar

2 w

ith

in

teg

ral

sh

ear

/ m

om

en

t att

ach

me

nt.

Ti S

tub

Rib

Al sp

ar

3

Al

/ T

i S

tub

sp

ar

4 w

ith

in

teg

ral

sh

ear

/ m

om

en

t att

ach

me

nt.

T

i S

tub

Rib

T

i T

ip R

ib

Al sp

ar

5

Al / T

i R

ib 3

Al

/ T

i S

tub

sp

ar

6 w

ith

in

teg

ral

sh

ear

/ m

om

en

t att

ach

me

nt.

Ti

Aft

sp

ar

wit

h

inte

gra

l

sh

ear

att

ach

men

t.

Actu

ato

r b

ay

Wh

eel

bay

T

i 4 p

art

Fw

d R

oo

t R

ib

Ti 2 p

art

Fw

d R

oo

t R

ib



184

Figure 137:- Final FB-24 / A-24 wing structural layout overview detailed in the

text below and in figure 140.

The final stage of wing maturation saw the deletion of the aluminium substructure

core to reduce weight for higher structural stiffness during relatively long periods of

high speed and high temperature flight and enhance damage resistance as shown in

figures 137 and 138. One concern the author had was the possibility differential

expansion between the aluminium and titanium components and the carbon BMI skin

with the aluminium likely to expand to an extent where cracking around fastener

sleeves could become an issue. Therefore the solution would be to move to an all

titanium substructure unlike that of the F/A-22A, (references 2, and 14), which

although originally planed as composite core wing had to retrospectively incorporate

titanium spars, this new wing would have no other substructure material. Although

more expensive to produce this would be traded – off over time by much higher

durability of the structure in terms of fatigue and thermal stress resistance, and

titanium has been proven to be fully compatible with carbon BMI skins in high

temperature sustained supercruise conditions by the F/A-22A EMD test phase, weight

would also be traded – off with reductions in the thickness of structural members.

In this final phase the undercarriage attachment and support structure was designed

although only in initial form and was quite possibly over engineered as time did not

permit the structural analysis the author had intended to undertake and a

rationalisation of this structure is proposed for further work undoubtedly it could

withstand the weight of the FB-24 or A-24 but could also by employed on an SR-71.

That having been said the undercarriage loads are carried out of the lugs on the two

ribs to the support stub spar and the rear torsion box trailing edge spar which appears

to be a satisfactory load distribution. Note although the flaperon actuator bay appears

shallow it is sited directly above the wing root faring which offers 20cm3 of growth if

Ti structural members

Undercarriage attachment lugs

Moment fittings

Fwd shear fitting

Aft shear fitting



185

required by extending out to the lower fairing skin without resorting to additional

blister fairings as in the case of the F-35 an F/A-22A. As will be described in the next

subsection the skins extend inboard of the fairing so that the root rib fastener line is

not exposed, on the top. This blending accommodates both RCS and aerodynamic

requirements without compromising either.

Fig

ure

138:-

Det

ail

ed d

escr

ipti

on

of

fin

al

win

g t

ors

ion

box s

tru

ctu

ral

layou

t.

Ti S

tub

Rib

Ti F

lap

ero

n a

ttac

hm

en

t R

ib 1

Ti F

lap

ero

n a

ttach

men

t R

ib 2

Ti S

tub

Rib

T

i T

ip R

ib

Ti K

ick

ed

Sp

ar

2

Ti S

tub

Sp

ar

1

Lead

ing

ed

ge T

i S

pa

r

Ti S

tub

Sp

ar

3

Ti K

ick

ed

Sp

ar

4

Ti In

term

ed

iate

Sp

ar

5

Ti U

nd

erc

arr

iag

e

att

ach

men

t S

par

5

Ti F

lap

ero

n a

ttach

men

t R

ib 3

Tra

ilin

g e

dg

e T

i S

par

Fla

pero

n a

ctu

ato

r b

ay

Fla

p a

ctu

ato

r b

ay

Un

derc

arr

iag

e a

ctu

ato

r b

ay

T

i R

oo

t R

ib

Un

derc

arr

iag

e a

ttach

men

t lu

gs a

nd

su

pp

ort

stu

b r

ib a

nd

stu

b s

par

str

uctu

re Ti S

tub

Sp

ar

6



186

4.5.2:- Fuselage layout and propulsion and weapons bay integration.

4.5.2.1:- Forward fuselage structure.

The forward fuselage retains the same basic OML shape as the F-35C including the

DSI intake however internally the structure was stiffened to accommodate the cut-out

for the second crew member as shown in figure 139, and this was aided by adopting a

single nose wheel undercarriage for land based operations rather than the larger twin

wheel catapult launch capable / carrier landing nose undercarriage of the F-35C. This

permitted substantially larger continuous keel beams, which do not have to diverge

around a broad wheel bay as shown in figure 140. The chine longerons extend the full

length of the forward fuselage sweeping upward at the intake to become thickened

shoulder longerons either side of the second crew station and additional intake chine

longerons were added outboard of the intake ducts. More substantial canopy landing

longerons were also incorporated for the aircraft commander (pilot) station with the

option of extending these further aft to form a hatch frame landing for the second

crew position, as currently this was to be incorporated into the forward fuselage top

skin. The radar array provision was also doubled over that of the F-35C by

incorporating a chevron mount swept at 520 to match the wing sweep with the radome

mounting line, giving a practical constant field of regard of 2400 which was intended

to resolve one of the issues raised in figure 97.

Initially high speed machined aluminium alloy was considered for the forward

fuselage substructure however issues of weight and thermal expansion differences

within the carbon BMI skin so alternatives were reviewed and at an early stage even

titanium was considered as in the Mach 3 SR-71. However reference chapter 4

revealed polyamide resin based carbon composites of the PMR family to the author,

which are compatible with carbon BMI up to temperatures of 1700 over long exposure

times, and have been used on advanced military aircraft substructure. Although these

structures have been produced by the pre-preg route which is a costly manufacturing

method, efforts are being made to develop a resin transfer moulding process for this

material by reducing its viscosity, by dissolving the resin in solvent, or in a low

viscosity reactive polymer, or chemical engineering to introduce twists into the

backbone of the polymer, and the author envisages an RTM route being available

early in the design cycle of the FB-24 / A-24 airframe. The major advantages of this

material are stability at high temperatures high stiffness and resistance to most

chemicals; however the cost of the forward fuselage structure would increase

substantially if RTM failed to mature as the main processing route for this material.

The principle dimensions of crew station size and longeron separation in this structure

are shown in figure 141 below, and the largely identical structure for the A-24 is

shown in figure 149 below.

4.5.2.2:- Centre fuselage structure.

The centre fuselage being less susceptible to kinetic heating than the forward fuselage

or the wings could be produced using high speed machined aluminium for the

substructure with the exception of the last frame which was the engine trunnion

attachment frame and must withstand close proximity to the higher temperature

section of the engine than the frames ahead of this engine mount frame. This frame



187

was also the mate joint frame for the aft fuselage and the wing torsion box trailing

edge spar attachment frame.

Fig

ure

139:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f fr

on

t fu

sela

ge

of

the

FB

-24 t

op

vie

w.

Ch

evro

n A

ES

A r

ad

ar

mo

un

t fr

am

e t

ota

l

mo

un

tin

g a

rea =

1.2

7m

Sta

rbo

ard

Fo

rwa

rd

ch

ine lo

ng

ero

n

Sta

rbo

ard

In

tak

e

ch

ine lo

ng

ero

n

Sta

rbo

ard

Fo

rwa

rd f

usela

ge

an

d h

atc

h s

up

po

rt lo

ng

ero

n

Sta

rbo

ard

Can

op

y

su

pp

ort

lo

ng

ero

n

Po

rt F

orw

ard

ch

ine

lon

gero

n

Po

rt C

an

op

y

su

pp

ort

lo

ng

ero

n

Po

rt F

orw

ard

fu

sela

ge

an

d

hatc

h s

up

po

rt lo

ng

ero

n

Po

rt In

tak

e c

hin

e

lon

gero

n

MO

SC

hatc

h h

ing

e f

ram

e

MO

SC

hatc

h f

ram

e s

eg

men

ts

Inta

ke s

up

po

rt a

nd

att

ach

men

t fr

am

es

N.B

.:-

All s

tru

ctu

ral m

em

be

rs a

re

RT

M C

arb

on

PM

R-1

5



188

Fig

ure

140:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f fr

on

t fu

sela

ge

of

the

FB

-24 u

nd

ersi

de

vie

w.

Bu

ild

jo

int

fram

e

582.0

mm

Tw

in c

on

tin

uo

us k

eel

bea

ms

No

se w

hee

l b

ay

308.8

mm

No

se w

hee

l att

ach

fra

me

Ch

evro

n A

ES

A r

ad

ar

mo

un

t fr

am

e L

O c

oate

d

field

of

reg

ard

240

0

Bif

urc

ate

d in

take d

ucti

ng

4.8

m

0.3

5m

N.B

.:-

All s

tru

ctu

ral m

em

be

rs a

re

RT

M C

arb

on

PM

R-1

5



189

Figure 141: - Principle structural dimensions for the spacing of the longerons,

and the nose wheel bay. Also shown are the dimensions of the MOSC crew hatch.

The angle of the chevron frame mounted AESA radar is also shown (the

mounting area of 1.27m2 compared to 0.682m

2 for the conventional F-35C

mount).

All of the wing attachment lugs interface onto titanium clevis lugs attached to the

aluminium frames (which there was not time to model) and all of the aluminium

structure has class packers where it interfaces with the carbon BMI skin to prevent

galvanic corrosion. Two substantial full length longerons and two full length keel

beams resist fuselage bending loads, and the weapon bays internal walls form

additional partial length keel beams, as shown in figures 142 through 144 below. The

frames consist of a major load baring core frame for the first four frames being closed

out by the weapons bay, and by fairing close out stub frames on the rear four frames

as shown in figure 145 (a), which differs considerably from the F-35 centre fuselage

frames as shown in figure 145 (b), where all three segments were primary load

bearing structure and consequently had to be of heaver construction.

4.5.2.3:- Aft fuselage structure.

The aft fuselage structure was all titanium from the outset and the approach was to

make the structure as ridged as possible, because of the two booms which house the

ruddervator actuators, and will move relative to each other and to this end a large

number of frames longerons and keel beams are employed. This structure was also

intended for structural analysis which is highlighted for further work, to determine if

additional keel beams are required in the F-5R and F-5L tanks running out to link

with the boom actuator mechanism beams which the author suspects or would the

20mm closure skins be capable of maintaining rigidity. The current structure is shown

in figures 146 and 147, the Niobium coated engine bay / tank nacelle skin was

3.4m

1.2m

1.5m

2.0m

520

0.9m

Nose wheel bay



190

replaced by carbon BMI on grounds of cost weight and effectiveness because none of

the surrounding structure could withstand a fire of 13700C should it occur. The land

based arrester hook is attached to the bottom of first frame of the aft fuselage, and the

rest of the frames split at the base to allow engine removal.

Fig

ure

142:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f ce

ntr

e fu

sela

ge

of

the

FB

-24 t

op

vie

w.

En

gin

e a

ttach

men

t fr

am

es A

l fo

rward

of

max t

em

pera

ture

zo

ne a

nd

Ti at

tru

nn

ion

att

ach

men

t / m

ate

fra

me

Win

g a

ttach

men

t fr

am

es A

l ex

cep

t fo

r re

ar

fram

e w

hic

h i

s

Ti an

d is t

he e

ng

ine / b

uil

d m

ate

jo

int

fram

e.

Al fu

ll len

gth

co

nti

nu

ou

s s

ecti

on

lon

gero

ns 7

0m

m b

y 5

0m

m

Ti

sid

e

walled

w

eap

on

s

bays

att

ach

ed

to

th

e

fram

es

form

ad

dit

ion

al lo

ad

path

En

gin

e t

run

nio

n

att

ach

men

ts

En

gin

e lu

g a

tta

ch

men

t

Un

derc

arr

iag

e lo

ad

carr

yin

g f

ram

es

So

ft b

uild

jo

int

1.7

9m

1.3

4m

2.2

4m



191

0.8

5m

Weap

on

s b

ay w

alls f

orm

ad

dit

ion

al lo

ad

path

s

Tw

o c

on

sta

nt

0.3

2m

dep

th K

icked

ke

els

So

ft b

uild

jo

int

Fig

ure

143:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f ce

ntr

e fu

sela

ge

of

the

FB

-24 u

nd

ersi

de

vie

w.

0.4

7m

7.6

m



192

Figure 144:- FB-24 / A-24 common centre fuselage structural arrangement plan

view, showing the propulsion wing integration.

Figure 145:- Comparison of the FB-24 / A-24 common centre fuselage structural

frame philosophy (a), with the F-35 frame philosophy of the 230-5 OML

configuration (b).

The aft fuselage bending loads are resisted by 100mm deep lower longerons and

70mmm by 50mm shoulder longerons as shown below. The ruddervator spigots are

Shear attachments

Moment attachments

Undercarriage load

carrying frames

Core fuselage

frame

Close out frames Split frames

FB-24 / A-24

Wing spar /frame

F-35

Wing joint

Shear attachments

(a) (b)



193

mounted through bearings in the boom spigot mounting plates, and are driven by the

actuator mechanism housed and mounted between the actuator drive bay keels shown

in figure 147 below. The structure of the ruddervators is shown below in figure 148

and is elaborated on there.

Fig

ure

146:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f aft

fu

sela

ge

of

the

FB

-24 t

op

vie

w.

Sta

rbo

ard

Ti fu

ll l

en

gth

co

nti

nu

ou

s s

ecti

on

sh

ou

lder

lon

ge

ron

s 7

0m

m w

ide b

y 5

0m

m d

eep

Sta

rbo

ard

Ti fu

ll l

en

gth

aft

fu

sela

ge c

hin

e lo

ng

ero

n

Po

rt T

i fu

ll l

en

gth

co

nti

nu

ou

s s

ecti

on

sh

ou

lder

lon

gero

ns 7

0m

m w

ide b

y 5

0m

m d

eep

Po

rt T

i fu

ll l

en

gth

aft

fu

se

lag

e c

hin

e lo

ng

ero

n

F-1

36-2

En

gin

e

Sta

rbo

ard

Ca

rbo

n B

MI

Nacell

e s

kin

Po

rt C

arb

on

BM

I

Nacell

e s

kin

Sta

rbo

ard

bo

om

actu

ato

r

sp

igo

t b

eari

ng

pla

te

Po

rt b

oo

m a

ctu

ato

r

sp

igo

t b

eari

ng

pla

te

N.B

.:-

All s

tru

ctu

ral

me

mb

ers

are

Ti

Ru

dd

erv

ato

r

Sp

igo

ts



194

Fig

ure

147:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f aft

fu

sela

ge

of

the

FB

-24 u

nd

ersi

de

vie

w.

N.B

.:-

All s

tru

ctu

ral

me

mb

ers

are

Ti

F-1

36-2

En

gin

e

Po

rt b

oo

m a

ctu

ato

r b

ay

s

Sta

rbo

ard

bo

om

actu

ato

r b

ays

Po

rt b

oo

m a

ctu

ato

r d

rive

bay k

eels

Sta

rbo

ard

bo

om

actu

ato

r

dri

ve b

ay k

eels

Po

rt C

arb

on

BM

I

Nacell

e s

kin

Sta

rbo

ard

Ti fu

ll l

en

gth

co

nti

nu

ou

s s

ecti

on

lo

wer

lon

gero

ns 3

0m

m w

ide b

y 1

00m

m d

eep

Po

rt T

i fu

ll l

en

gth

co

nti

nu

ou

s s

ecti

on

lo

we

r

lon

gero

ns 3

0m

m w

ide b

y 1

00m

m d

eep

Sta

rbo

ard

Ca

rbo

n

BM

I N

acell

e s

kin

F

R-5

Fu

el ta

nk b

ou

nd

ary

fra

mes

FL

-5 F

uel ta

nk b

ou

nd

ary

fra

me

s

Lan

d

based

a

rreste

r

ho

ok a

ttach

men

t fr

am

e



195

Figure 148:- Ruddervator internal structural layout (Port Ruddervator shown),

the ruddervator skin material would be a quartz single wave pass composite to

permit the radar emissions to enter the structural RAM core material beneath,

this structural RAM core would be used on all control surfaces.

Figure 149:- Final AI installation into completed structural forward fuselage

layout, structurally identical to the FB-24 layout only without second crewman

occupancy provision or fuselage hatch.

Unitised Ti substructure frame machining

from initial near net forging

Ti spigot mating fixture machining from

initial near net forging

Structural RAM honeycomb

filled with absorbent foam

Same structural members as forward fuse of FB-24 for manufacture and assembly commonality and cost reduction

All structural members

are RMT carbon PMR-15



196

Figure 150 above shows the full integration of the major systems with the major

structural build modules within the carbon BMI skin. Shown in figures 151 and 152 is

the F-35C internal structural layout for comparison with that of the FB-24.

Fig

ure

150:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f fi

nal

FB

-24 c

on

figu

rati

on

.

Fo

rward

fu

sela

ge b

uild

mo

du

le in

ca

rbo

n P

MR

-15

Cen

tre f

use

lag

e b

uil

d

mo

du

le in

Al

an

d T

i

Weap

on

s b

ay

s T

i S

PF

/DB

Carb

on

BM

I sk

in

Cera

mic

/ S

tru

ctu

ral R

AM

lead

ing

ed

ge

fla

p

Cera

mic

/ S

tru

ctu

ral R

AM

lead

ing

ed

ge

fla

p

Aft

fu

sela

ge b

uild

mo

du

le in

Ti

Cera

mic

/ S

tru

ctu

ral

RA

M f

lap

ero

n

Cera

mic

/ S

tru

ctu

ral

RA

M f

lap

ero

n

Win

g s

tru

ctu

re T

i

Win

g s

tru

ctu

re T

i

Ti / S

tru

ctu

ral R

AM

rud

derv

ato

rs

Ti

fire

sh

ield

n

ot

sh

ow

n

FS

10

149 –

to

–F

S15

893.



197

Figure 151: - F-35C side view internal structural layout 230-5 configuration.

Figure 152: - F-35C plan view internal structural layout 230-5 configuration.

The substructure of the FB-24 / A-24 common airframe has been described above to

the extent of its current maturation, and during the preparation of this work the F-35C

has also matured beyond the 230-5 configuration used above for comparison, but the

matured structure is ITAR restricted. Therefore when this work is used in the future

the reader is reminded that the capabilities quoted for the F-35C represent only the

230-5 configuration and not any final SDD or production aircraft. This concludes the

structural description section of this thesis, and general arrangement views with key

dimensions are presented in figures 156 through 158 below.



198

4.6 Fuel system integration and tank layout of the FB-24 and A-24.

The fuel tank layouts for the two crew FB-24 and the A-24 UCAV shown below in

figures 153 and 154 respectively the airframe contours and ruddervators are shown for

special orientation. The fuel tank system for both aircraft was identical except for the

provision of the F-1 tank in the forward fuselage of the A-24, which could also be

installed in a single seat version of the FB-24 if that were adopted as a direct F-35

replacement. The dimensions were determined from the measurement of CATIA V5

solid models produced by the author, and the capacities were calculated from the data

provided in reference 28and presented below: -

JP-5 (RAF) (JP-5 Mil – spec density of 6.82lb/US gal equal to 51.10lbs/ft3).

JP-8 (USAF) (JP-8 Mil - spec density of 6.80lb/US gal equal to 50.86lbs/ft3).

JP-4 (Special fuel) (JP-4 Mil – spec density is 6.55 lbs/US gal equal to 49lbs/ft3).

From these values and the volumes measured from the CATIA models the total fuel

capacity using JP-8 as the standard fuel selected for both aircraft:-

For the FB-24 in two crew configuration of 22,238lb reduced to 21,126lb with foam

filling:

The A-24 UCAV capacity of 26,458lb also reduced with foam filling to 25,135lb.

These values compare favourably with the standard F-35C which in 230-5 OML form

had a publicly quoted internal fuel capacity of 19,100lb of JP-8 and also indicate that

a single seat version of the FB-24 has the potential for substantially longer range

missions. The fuel tanks specified were of the self – sealing integral foam filled type,

common in modern combat aircraft which results in 2.5% of the fuel volume being

lost to displacement of fuel by the foam and a further 2.5% of the fuel being absorbed

by the foam. Consideration in the fuel tank layout was given to the method of in-flight

refuelling and the placement of the Integrated Power Pack Unit (IPPU) and the units

associated exhaust ducting.

Unlike the current USAF fighter inventory in-flight refuelling of this aircraft was to

be conducted using the „probe and drogue‟ method instead of the „flying boom‟ which

although representing a lower fuel transfer rate, would enable more aircraft to refuel

from a single tanker and commonality with the USN / USMC / and NATO allies who

all use the „probe and drogue‟ system. This system removes complexity from the

centre fuselage and enabled the enlargement of the F-2 and F-4 tanks and the resulting

increased internal fuel capacity. The provision of the systems envelope and the weight

balance required by the IPPU reduced the scope for extending the F-4 tank below the

level of the weapons bays which form the lower boundary of the F-2 tank in the fuel

system layout, because the IPPU location was a critical systems placement for the

following reasons: - structural support of the IPPU, systems routing, minimise exhaust

gas signature, and impingement of the exhaust gas on the left ruddervator skin, the

exhaust would be re-routed to exit lower left below the aircraft fuselage instead of

above it as in the F-35 family ( there was not time to model this and it fall into the

scope of future work). This complete the fuel system integration section of this

conceptual design study within the constraints of this thesis.



199

Figure 153:- The FB-24 fuel tank layout and sizing in volume and capacity with a

total capacity for JP-8 = 22,238lb equal to 21,126lb when the reduction for foam

filled tanks is made. Note airframe boundary and ruddervators for reference.

Figure 154:- The A-24 fuel tank layout and sizing in volume and capacity with a

total capacity for JP-8 = 26,458lb equal to 25,135lb when the reduction for foam

filled tanks is made. Note airframe boundary and ruddervators for reference.

F-3R tank volume = 4.1m3

capacity = 5,905lb JP-8 fuel

F-2 tank volume = 3.08m3


F-3L tank volume = 4.1m3

capacity = 5,905lb JP-8 fuel F-4 tank volume = 2.66m3


F-5R tank volume = 0.751 m3


F-5L tank volume = 0.751 m3


F-3L tank volume = 4.10m3




F-5L tank volume = 0.751 m3


F-5R tank volume = 0.751 m3

capacity = 1,081lb JP-8 fuel F-3L tank volume = 4.10m

3






IPPU Location

IPPU Location



200

Fig

ure

155:-

FB

-24 P

lan

vie

w G

ener

al

Arr

an

gem

ent

dra

win

g.

52

0

63

0

5.945m

15.3

20

1.4

0m

0.8

1m

0.6

1m



201

Fig

ure

156:-

FB

-24 S

ide

vie

w G

ener

al

Arra

ngem

ent

dra

win

g.

55

0

19.1

53m

3.7

90m

8.0

86m



202

Fig

ure

157:-

FB

-24 G

ener

al

Arra

ngem

ent

dra

win

g F

ron

t vie

w.

50

0

3.3

28m

5.9

45m

13.7

22m



203

5.0 Conclusions and further work recommendations.

The objective of this thesis was to produce a conceptual design study for a force

package consisting of a two seat advanced interdiction aircraft and a complementary

UCAV capable of replacing the UK Royal Air Force Tornado, Royal Australian Air

Force F-111, and USAF F-117 and F-15E air assets, both of these aircraft would have

greatly enhanced capabilities in stealth, range, and supercruise capability. The

performance being measured against a representative future mission profile produced

by the American Institute of Aeronautics and Astronautics, and the USAF Academy.

Table 13: - Comparison of RFP with F-24 / A-24 Capabilities.

Customer Needs RFP Requirements F-24 / A-24

Capabilities Meets Need

Maximum Mach number 45K ft

1.6 1.6 / 1.6 Yes

Weight GTOW 71,000lbs 53,187lbs / 56,897lbs Yes

Max Span Not > 65ft 45.02ft Yes

Accommodation 2 crew / AI unit 2 crew / AI unit Yes

Weapons internal load

4,614lbs 4,194lbs / 3,974lbs No*

Max Range internal fuel.

1800nm 1,283nm / 1,453nm No*

SEP, M=0.9, mil thrust, h=0.

0ft/s 200ft/s Yes

SEP, M=0.9, max thrust, h=0, n=1.

300ft/s 800ft/s Yes

SEP, M=0.9, h=0,n=5 max

thrust 50ft/s 670ft/s Yes

Maximum Instantaneous turn rate h=0

8.0 deg/s 18 deg/s Yes

Cost 500units $75,000,000.00 $ 45,823,547.00 Yes

*Capability can be met within weight limit and engine thrust limits and cost

limits.

The configuration design and parametric analysis using both classical analysis and the

AeroDYNAMIC V2.08 analysis tool set resulted in the analysis models being of a

much cruder level of definition, which made the drag results pessimistic because the

fuselage had to be defined as a ellipse and therefore the blending and side wall

sloping which reduced drag as well as RCS but this could not be evaluated. This

phase was originally intended to include CFD analysis of the CATIA V5 surface

models produced enabling precise drag analysis but this could not be incorporated

because of the time scales involved in the project. As a result the two best

configurations NB1 and NB2 went forward for analysis with AeroDYNAMIC V2.08.



204

The NB1 being a trapezoidal wing planform with a conventional empennage

configuration with two fixed vertical tails and rudders, and two all moving horizontal

tails. The NB2 configuration had the same trapezoidal wing and fuselage, but two all

moving ruddervators instead of the conventional empennage arrangement of NB1.

From the AeroDYNAMIC analysis there was a drag penalty for the conventional

tailed NB1 configuration which the author expected, and a weight penalty as can be

seen in table 11 in section 3 of the thesis, which contains weights and cost data for the

two configurations based on the RAND Corp DAPCOIV model within

AeroDYNAMIC. These results also indicate that both aircraft could be built below

the target values in the USAFA / AIAA Request For Proposals which is more a

reflection of the level of sophistication and fidelity of modelling tool and a much

more accurate tool is needed to determine if this assertion is close to reality. As a

result of completion of this phase of the design study the NB2 ruddervator

configuration was selected to be matured into a full structural concept layout in phase

two, the NB1 concept configuration was held in reserve and not developed further.

The major structural component layout of the complete aircraft and systems

integration, without detailed structural sizing of the substructure beyond notional sizes

based on legacy aircraft and academic texts due to time constraints. The primary

objective was to determine a structural layout capable of accommodating the principal

systems i.e. the aircrew, AI (UCAV) system, propulsion system, weapons system, and

fuel system. The structural arrangement contained within section 4 of this thesis

accomplishes this objective incorporating advanced high temperature composite

materials namely carbon BMI for all aircraft skins and carbon PRM-15 for the

forward fuselage, high speed machined aluminium, and titanium for the centre

fuselage and wing structure, and titanium exclusively for the rear fuselage, with

ceramic skinned structural RAM used for the control surfaces and high lift devices.

The weapons bay size was increased over that of the current F-35C aircraft to

accommodate JASSM internally as well as all of the weapons contained in the

Request for Proposals. The aircrafts ability to supercruise will be dependent on the

development of the F-136-2 YF-120 derived engine, however the two crew FB-24

aircraft has a fuel capacity of 21,126lb and the unmanned A-24 has a fuel capacity of

25.135lb both exceeding the current F-35C value of 19,100lb. The current F-35 radar

frame was replaced by the author‟s chevron design swept at the same angle as the

wings this doubles the area available for AESA array mounting and substantially

increases the field of regard, and this could be incorporated into the current F-35

airframe.

The final FB-24 the manned and A-24 unmanned Advanced Interdiction Aircraft

contained herein has common forward fuselage skin OML with the F-35 aircraft in

order to retain a degree of commonality, the rear fuselage top skin OML is also

similar. The build modules are different consisting of three fuselage modules instead

of four for the F-35, this new arrangement being more structurally efficient, and the

wing attachment philosophy is the conventional fuselage side interface with a two

piece wing instead of the current continuous wing with split frames of the F-35.

Although the FB-24 and A-24 airframes evolved into a larger configuration and differ

in structural materials and layout from the F-35C they can be seen as a next

generation long range Joint Strike Fighter and as a follow on in the same way as the

F/A-18E and F have grown out of the F/A-18 Hornet program, and these aircraft

would still be less expensive than a clean sheet of paper design. Although at their



205

current weights they are below the RFP targets in payload and range there is more

than sufficient weight growth margins for increased fuel and payload capabilities.

Future work required to move this design proposal forward into a preliminary design

is outlined below: -

1. Detailed parametric analysis, using a more sophisticated analysis tool.

2. Computational Fluid Dynamics investigation of the true OML drag values

through the following Mach numbers: - 0.8: 0.9:1.0:1.4:1.5:1.6 and 1.8 as well

as 1800 turns at Mach 1.6 with control surfaces deployed at angles throughout

their range of motion given in section 3, to determine control effectiveness and

drag contribution.

3. Optimisation of the structural layout by using hand calculations and the results

of the analysis in further work activity 1 to create a Vn diagram for each major

build component, and then create a Finite Element model of the structure

based on the initial layout models created by the author who would be willing

to supply them through Cranfield University School of Engineering. This

analysis would start with the wing structure to rationalise the main

undercarriage interface, and then explore the kicked spare to rib interface

joints and the control surface and high lift device interfaces, and thereafter

develop a skin model to determine pad up requirements, then moving on to

address the issues raised in section 4 with respect to the stiffness of the rear

fuselage booms.

4. Develop a conceptual design for the MOSC hatch and determine the need for

any additional landing support longerons within the forward fuselage

structure, over the provisions which could be made in the skin.

5. Model the wing to centre fuselage to wing interface joints on the frame side

using the wing fittings as a base point, to determine if the aluminium frames

could take the wing interface loads distributed through Ti bolt on fittings as

the author feel would be possible.

6. Determine more accurate component weight values for adoption as targets for

the as draw structure.

These are the main issues which need to be addressed before PDR can be

contemplated and could form the basis for another Airframe Engineering students

study at Cranfield University.

6.0 References and literature review:-

1) (a) Aircraft Engineering Projects for Engineering Students: by Jenkins.

L. et all: Published by the AIAA Education Series 2003, as referenced

in section 1 introduction and 2 requirements capture.



206

1) (b) Aerodynamic V3.0 software produced by the USAFA and released

through the AIAA Education Series 2004, as referenced in sections 1,

introduction and 2 requirements capture.

2) F-22 Raptor pages 87-88: by Sweetman. B.: Published by MBI in the

USA 1998, as referenced in section 2 requirements capture.

3) Iron Hand (Smashing the Enemy‟s Air Defences), pages 284-285:

Published by Patrick Stephens Ltd 2002, as referenced in section 2

requirements capture.

4) Aircraft Concept Design Synthesis page 7: by Howe. D.: Published by

Professional Engineering Publishing 2000, as referenced in section 3

FB-24 and A-24 concept design and selection.

5) Fundamentals of fighter design 1st edition: Whitford R: Published by

Airlife 2003, as referenced in section 3.

6) Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by

Harkins, H.: Published by Centurion Publishing UK 2004, as

referenced in section 3.

7) Aerodynamics for Engineers 4th

Edition USAFA: by Bertin. J. J:

Published by Prentice Hall 2002, as referenced in section 3.

8) “Wind - Tunnel Investigations of Variable Camber and Twist Wing”

TND-8457: NASA: Aug 1977, as referenced in section 3.

9) Lockheed Martin F/A-22A Raptor: Stealth Fighter: by Miller. J.

Published by Midland Counties Publishing 2005.

10) Evaluation of Turbulent Models for High – Lift Military Airfoil

Flowfields: by Kern. S: AIAA96-0057, as referenced in section 3.

11) NASA Dryden Fact Sheet –F-16XL-1 Testbed aircraft as referenced in

section 3.

12) www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html, as


13) Lockheed Martin‟s Affordable Stealth paper by Haisty B.S, Published

by Lockheed Martin Aeronautics Washington D.C. 2000, for the

National Press Club.

14) F-22 Raptor (Americas next lethal war machine) by Pace. Steve:

Volume 1 of Walter J. Boyne Military Series. Published by McGraw

Hill 1999, as referenced in section 3.

15) The F-22 web site www.F-22.com.html as referenced in section 3.

http://www.dftc.nasa.gov/Newsroom/FactSheets/FS-051-051-DFRC.html

http://www.f-22.com.html/



207

16) The AFT Contenders YF-22 and YF-23 Air Superiority into the 21st

Century 3rd

edition. By Sun Andy: Published by Concord Publications

Ltd. Hong Kong 1991, as referenced in section 3.

17) Superfighters (The next generation of combat aircraft) by Williams

Mel. Published by Airtime publications 2002.

18) Aircraft Performance and Design: by Dr Anderson. D. John. Jr

University of Maryland: Published by McGraw Hill 1999, as


19) Modern Combat Aircraft Design: by Huenecke Klaus: Published by

Airlife Publishing 1987, as referenced in section 3.

20) Lecture notes AERO-481 Lesson 12: Survivability Propulsion

Integration and Systems, USAFA on reference 1 software title, as


21) Aircraft Design A Conceptual Approach 3rd

Edition: by Raymer,

Daniel. P.: Published by the AIAA 1999, as referenced in section 3.

22) Introduction to Aeronautics a Design Perspective: by Brandt S. A. et

al: Published by AIAA 1997, as referenced in section 3.

23) Jet Bombers (From the Messerschmitt 262 to the Stealth B-2): by

Gunston B. et al: Published by Osprey Aerospace 1993, as referenced

in section 3.

24) Modern Fighters: by Spick. M.: Published by Salamander Books ltd

2000, as referenced in section 3.

25) Fundamentals of Fighter Design 2nd

Edition: by Whitford R.:

Published by Airlife Publishing 2005, as referenced in section 3.

26) Advanced Tactical Fighter to F-22 Raptor Origins of the 21st Century

Air Dominance Fighter: by Aronstein. D. C. et al Published by AIAA

1998, as referenced in section 3 and appendices B.

27) Have Blue and the F-117A evolution of the Stealth Fighter: by

Aronstein. D. C. et al: Published by AIAA 1997, as referenced in

section 3.

28) Airframe Structural Design (Practical design information and data on

aircraft structure): by Nui. M.: Published by Hong Kong Conmilit

Press Ltd 1997, as referenced in section 4.

29) Ultimate Fighter: the Lockheed Martin F-35 Joint Strike Fighter: by

Sweetman. Bill. : Published by Zenith Press 2004, as referenced in

section 4.



208

30) Marguerite Ozburn Global Strike Systems, Boeing IDS Business

Support, Communications and Community Affairs, P.O. Box 516 St

Louis MO 63166 e-mail) as referenced in section 4.

31) Lessons Learned from the F/A-22 and F/A-18E Development

Programs RAND_MG276: by Younossi. Obaid: Stem. David. E. et al:

Prepared by for the USAF by RAND Project Airforce, as referenced in

section 4.

32) Composite Materials for Aircraft Structures second edition: by Alan

Baker et al: Published by the American Institute of Aeronautics and

Astronautics 2004, as referenced in section 4.

33) Aeronautical Engineers Data Book: by Clifford Matthews: Published

by Butterworth Heinemann 2002, as reference in appendices E.



209

Appendix:

Appendices A: - Details of primary air and surface threats.

For this study the primary air to air threats are anticipated to be MiG-29C‟s:

MiG-31M‟s: and control aircraft of the A-50 Mainstay type because in the future

Global Strike ConOps force in which this aircraft is intended to operate

Eurofighter Typhoon, and F/A-22 Raptor air dominance fighters will be

responsible for tying down the Su-27C / P, and Su-30A enemy air assets, when

encountered. Also with the latter aircraft being the most expensive Russian

export platforms nations capable of procuring them are most likely to employ

them for free ranging attack fighters rather than homeland defence tied combat

air patrols. Although more capable point defence fighters such as the MiG Light

Weight Fighter Project figure A1 or the Chinese J-10 figure A2, could pose long

term treats to both the FB-24 and A-24 components of the mixed fleet, however

the J-10 could be detected and avoided and the Russian LWFP may never

appear as a real aircraft.

The MAPO MiG-29M: air defence fighter.

The statistics of this aircraft from published data (References 8:-Pages 60-76: Modern

Fighters – The Ultimate Guide to In Flight Tactics, Technology, Weapons and

Equipment: by Spick Mike: Published by Salamander books ltd London 2000 and

Reference 9:- Pages 15-67: Russian Air Power – 21st Century Aircraft, Weapons and

Strategy: by Gordon Yefim and Dawes Alan: Published by Airlife Publishing Ltd

2002) are given below:-

Dimensions: - Span = 37ft 3in: Length = 57ft 0in: Height = 15ft 6in: Wing Area =

409ft2

Aspect Ratio: - b2/S = 3.40

Weight: - Empty WE = 24,250lbs: Take off WTO = 37,038lbs

Power: - 2 x Klimov RD – 33K engines delivering a maximum installed thrust of

19,400lbs, and a military thrust of 12,125lbs each.

Fuel: - Internal = 9,978lbs: fuel fraction = 0.27

Loading: - Thrust to weight (maximum) = 1.05: Wing Loading = 91lbs/ft2

Performance: - Vmax high = Mach 2.30: Vmax low = Mach 1.23: Vmin = 110kt:

Operational Ceiling = 55,777ft: Climb Rate = 64,945ft/min

Weapons: - One Gryazev / Shipunov GSh 301, 30mm cannon with 100 rounds:

Two R-27R (AA-10a Alamo-A) medium range active radar homing missiles: and

up to Six R-73 (AA-11 Archer) short range infra – red homing missiles or Six R-

60M (AA-8 Aphid) short range infra – red homing missiles.



210

Figure A.1:- MiG LWFP proposal highly manoeuvrable light weight fighter to

rival the F-35 and defend against UCAV‟s, a questionable project, source

Aviation Week & Space Technology (AW&ST) August 2003.

Figure A.2:- The Chinese J-10 comparable with the BAE Systems / SAAB

Gripen, this late 4th

generation aircraft could be heavily exported to potential

enemy combatants over the next two decades because it is projected to be

cheaper than Russian MiG‟s but have more modern systems however it has no

stealth features and could be avoided, source AW&ST January 2006.



211

Air to Air missile capabilities: -

R-27R (AA-10a Alamo-A) intended to intercept and destroy aircraft, helicopters,

UAV‟s, UCAV‟s and ALCM‟s at medium ranges, day and night in all weather

conditions, from any angle, in ground or sea clutter and in the presence of a wide

range of countermeasures and defensive actions by a target.

The missile can attack targets at heights between 66ft and 88,583ft, and at speeds of

up to 1,890kts regardless of its initial position, within a field of view of +/- 50o. The

launch aircraft can pull 5G at the moment of missile release and the maximum height

differential between the launch aircraft and the target is 32,808ft above or below. For

the MiG-29 the AA-10 Alamo is mounted on two special underwing pylons.

R-73 (AA-11 Archer) short range heat seeking missile fitted with the Mayak (Beacon)

infrared detector head, is intended to intercept and destroy highly manoeuvrable

piloted and unmanned air vehicles, day and night from any angle in the forward and

rear hemispheres, in ground clutter and in the presence of intense ECM activity. This

missile has outstanding agility with the capability to destroy targets manoeuvring at

up to 12G. This missile can attack targets flying at heights between 66ft and 65,617ft

and speeds up to 1,350kts, from any initial position within a field of view of +/- 45o

and with an angular velocity of up to 60o/sec off boresight. Target designation for the

Archer‟s missile seeker head can be accomplished through the MiG-29‟s helmet –

mounted sighting and targeting system. All variants of the MiG-29 can employ the

AA-11 Archer missile which can be mounted on any of six underwing hard points on

the aircraft.

R-60M (AA-8 Aphid) this is an earlier short range heat seeking missile fitted with the

Komar-M (Mosquito) seeker head, and like the Archer it is intended for the

destruction of highly manoeuvrable targets but is limited to those which are within

visual range of the launch aircraft. The low launch weight of the missile and its

advanced aerodynamic layout give the Aphid outstanding agility and the capability to

destroy targets manoeuvring at up to 12G. This missile can engage targets flying at

heights between 98ft and 65,617ft and speeds of 1,350kts from any angle within the

pilot‟s field of view +/- 20o and at angular speeds of up to 35

o/sec off boresight. Like

the AA-11 the Aphid can be used for forward hemisphere engagements at short range

and the launch aircraft can pull 7G at launch without compromising the engagement

profile. The AA-8 Aphid has no equal in terms of weight and size in the west.

The MAPO MiG-31B: interceptor fighter.

The statistics of this aircraft from published data (References 8:-Pages 60-76: Modern

Fighters – The Ultimate Guide to In Flight Tactics, Technology, Weapons and

Equipment: by Spick Mike: Published by Salamander books ltd London 2000 and

Reference 9:- Pages 15-67: Russian Air Power – 21st Century Aircraft, Weapons and

Strategy: by Gordon Yefim and Dawes Alan: Published by Airlife Publishing Ltd

2002) are given below:-



212

Dimensions: - Span = 44ft 2in: Length overall = 74ft 5in: Height = 20ft 2in: Wing

area = 663ft2

Aspect Ratio: - b2/S = 2.94

Weight: - Empty WE = 48,104lbs: Take off WTO = 90,389lbs

Fuel: - Internal = 36,045lbs: fuel fraction = 0.40

Power: - 2 x Soloviev D-30F6 afterburning bypass turbojet engines delivering a

maximum installed thrust of 38,580lbs and a military installed thrust of

20,944lbs each.

Loading: - Thrust to weight (maximum) = 0.85: Wing loading = 136lb/ft2

Performance: - Vmax high = Mach 2.83: Vmax low = Mach 1.23: Vmin = 140kt:

Operational Ceiling = 67,589ft: Climb Rate = 41,000ft / min

Weapons: - One Gryazev / Shipunov GSh 6-23, 23mm six barrelled cannon with

260 rounds: Four R-33 (AA-9 Amos) long range (162nm) active radar homing

missiles developed exclusively for the MiG-31: or Four R-40 (AA-6 Acrid)

medium range active radar homing missiles: and Two R-60 (AA-8 Aphid) short

range infra-red homing missiles. Additional configurations offered include up to

Six R-37 advanced long range active radar homing missile, and Four advanced

R-77 (AA-12 Adder) active radar homing missiles capable of sustaining 12G

while manoeuvring and having a range of 54nm under the wings on individual

pylons in overload condition.

All variants of the MiG-31 possess the capability of countering mass air attacks by

both manned aircraft and cruise missiles, by being able to track several targets

simultaneously and to launch all their missiles against individually identified targets

in salvo. Four MiG-31‟s are for example, capable of covering a 540nm linear front,

with each aircraft simultaneously tracking ten targets and able to engage four. Unlike

current Western fighters which have to engage their targets in comparatively narrow

azimuth and range limits, the MiG-31 has an engagement zone of +/- 70o of boresight.

The Ilyushin / Beriev A-50 Mainstay AWAC aircraft

The A-50 differs from the basic transport Il-76 Candid, from which it was derived, by

having a slightly lengthened fuselage and more noticeably, by the presence of the 29ft

6in diameter rotodome of the Shmel (Bumblebee) AWAC suite (known by NATO as

Squash Dome). The usual operating height of the A-50 is around 32,808ft and the

normal operating endurance, without air – to – air refuelling AAR is four hours.

The A-50 carries a fifteen – man crew, of whom five are fight – deck crew (pilot, co-

pilot, two navigators and a flight engineer), plus a rear crew of ten to operate the radar

consoles and other mission equipment.



213

The entire mission complex of the A-50 is called the Shmel, developed by NPO

Vega-M, and includes the radar itself, Identification Friend or Foe IFF, the

information processing system and operators‟ consoles and an Electronic

Countermeasures ECM resistant digital communications suite, linking the A-50 with

both ground command posts, SAM‟s, ships, and fighter aircraft control networks. The

coherent pulse Doppler 3-D radar has a 360o circular scan and permits detection and

tracking of airborne targets including low flying cruise missiles against local terrain

(including water and any other type of surface relief, such as steppe, forest or

mountain).

The system is also capable of detecting and plotting surface shipping. Airborne –

controlled intercept (ACI) are achieved both by use of automated data – links and in

manual mode, using normal radio channels and plain – voice radio commands.

Information on the movements of targets of interest is transmitted to the Command

Posts of the Automatic Control System (KP ASU / Komandnyy Punkt

Avtomatizirovannykh Sistyem Upravleniya) of Russia‟s armed forces by digital data-

link through special relay centres. The transmission range of such information to the

KP ASUs is 189nm to 1,080nm.

When working at extreme ranges from any of the command posts satellite

communications links are used to transmit target data where available. Although the

threat nations may not have the satellite coverage capability a lower specification

Automatic Control System is offered with the A-50 package as well as French system

upgrades. The A-50 has colour displays which give alphanumeric data, including

track number, course, altitude, and speed of friendly interceptors, plus their remaining

fuel state, in panoramic format. Data can also be recorded in a documented text

format for onward transmission.

The aircraft is capable of AAR and is equipped with chaff and flare dispensers for self

protection. This asset will be heavily screened by three MiG-29 aircraft in the

immediate kill zone following Russian defence training patterns therefore the AIA

signature must be maintained in all flight phases especially manoeuvre and weapons

release of either Advanced Anti Radiation Weapon or the ALOSNW stores.

Development of the A-50 commenced in 1965, but the first aircraft only entered

service 1984 and by 1992 there were around twenty five in service in total, although

with the end of the Cold War only ten are in front line units in Russia, the remainder

are possible exports. Licensed production is being sought by the PRC and India, also

Iran with Russian assistance has produced an indigenous derivative of the A-50 from

stock Il-76 airframes, and North Korea intends to purchase surplus Russia A-50‟s

through an export company. (Reference10: -Combat Aircraft Magazine issues: -

Volume 3 No4 and Volume 4 No 5 Published by Ian Allan 2001 and 2002

respectively).



214

Primary surface to air threats come from the mobile S-300V SAM system, and

the highly mobile Buk-M1 (SA-11) replacement for the Kub (SA-6) SAM that

shot down Captain Scott O‟Grady‟s F-16 over Bosnia.

The S-300V (SA-12): - Gladiator SAM system.

The S-300V (SA-12a Gladiator) system has two missiles (developed from the anti-

ballistic missile SA-12B Giant SAM), one large and one small, the smaller one has a

peak velocity of Mach 6 and can destroy targets evading at 8g through clutter and

ECM / decoy systems over an effective range of 30nm at altitudes between 2,000ft

and 60,000ft, the second larger missile attains a peak velocity of Mach 8 and through

advanced terminal aerodynamics can destroy targets manoeuvring at 12g through

ECM / decoy systems at altitudes between 12,000ft and 80,000ft. Targeting

information can be obtained from the main Almaz NPO family of sensors or off bored

from the A-50 Mainstay or MiG-31 Foxhound airborne platforms, and are capable of

intercepting HARM missiles launched against them.

The on – board initial long range tracking and target acquisition radar system

currently exported is the frequency – agile F-Band Bill Board radar, which is backed

up by the Grill Pan phased array for aircraft tracking and launch / interception

solution generation, which is able to track up to twelve targets and control six SAM‟s

against them simultaneously. The S-300V is a mature system and has been fully

operational in Russia since 1986.

The S-300PMU-1 and the longer range S-300PMU-2 (SA-10 Favorit) variants are

specifically designed for the export market and are at time of writing operational in

North Korea, the Balkans, the Indo-Pakistan border area and Syria-all potential threat

regions.

In addition the Almaz S-400 Triumf (SA-20 Gargoyle) family which are improved

faster derivatives of the S-300V‟s are now available on the world market after

completing field trials. These use the same GAZ-66 and MAZ-543 6x6 or 8x8 chunky

wheeled self propelled launchers which can disperse from the central radar nervous

system along with survey, mess, dormitory and power generator modules fielded on

MAZ-543 chassis vehicles, allowing them to pop up at unexpected locations, and all

have A-50 and MiG-31 target data links. (Reference 12:- pages 284-285, Iron Hand

Smashing the Enemy’s Air Defences: by Thornborough. M. A. and Mormillo. B. F.:

Published by Patrick Stephens Limited an imprint of Haynes Publishing UK 2002)

The Buk-M1 (SA-11):- SAM system.

The Buk-M1 is the replacement for the Kub (SA-6) and is in full scale production for

the export market, with a single system which is mounted on 11 vehicles a defender

has 36 missiles ready to launch at any one time. The missiles have a reach of some 45

to 105nm, depending on model similar to that of the SA-10 Grumble, with a closing

speed of up to 14,770ft/sec. Also like the SA-10 this weapon system relies on radar

guidance coming from associated F-Band Continuous Wave pulse – Doppler Clam

Shell, tower - mounted Big Bird or 3-D Tombstone long – range surveillance / EW

radars, with I/J-Band Flap Lid phased array radar used for target tracking.



215

As with the S-300V and S-400 this system can be linked to the A-50 so that target

information can be obtained without the risk of the system being exposed to attack

from anti – radiation missiles.

Short – range Man portable SAM‟s.

The man portable SAM‟s MANPADS of the SA-14/-15/-16 have imaging infra-red

and ultraviolet seekers operating at both ends of the visible spectrum, and the SA-16

Gimlet simply ignores flares altogether, and over 1 million have been sold to date

most to threat nations. The best counter to these is to fly above 20,000ft to avoid them

and anti aircraft artillery (triple –A) fire, and rely on older precision guided munitions

PGM‟s or newer JDAM‟s and J-series weapons in a fast pass lob well away from the

target area. But such tactics place non-stealthy aircraft within the prospective shooting

range of the SA-11 SA-10, and S-300V and S-400 SAM‟s detailed above. (Ref 12)

Commenting on this surface to air threat environment one pilot put it, “You have to

hide from them until you can kill them”. New fourth generation stealth fighters such

as the F-35 Joint Strike Fighter and F/A-22 Raptor, relying on golf ball and marble

Radar Cross Section‟s (RCS) respectively, aim to do precisely that – hide, not evade,

and right over the enemy‟s noses. High – altitude attacks at up-to 40,000ft would

help them avoid the operational envelope of infra-red seeking SAM‟s and triple

A, and speeds of Mach 1.3 – to – Mach 1.5 would reduce the effective envelope of

the large radar guided SAM‟s like the S-300V‟s, S-400‟s, and SA-11‟s.(Ref 11

and 12), which is the most important factor in determining the mission profile of

the FB-24 AIA, and hence the aircraft itself.

Appendices B: - Signature control.

B.1 Radar Signature Control.

The threat to aircraft of radar detection and tracking comes from a variety of sources,

as shown in Table 6. Since radar range is a function of the fourth root of the radar

cross-section (i.e. RCS1/4

) an order of magnitude reduction in RCS, for example will

give a 44% reduction in detection range:

R1 / R2 = [RCS1 / RCS2]1/4

= [1 / 10]1/4

= 0.56

Similarly, the search area of the radar will be reduced to 32% and search volume to

18%. Hence a very large reduction in RCS not one tenth but one thousandth is

necessary to achieve a tactically useful effect (e.g. a reduction of 82% in an enemy

systems useful detection envelope).

RCS depends on the following features: - aircraft shape: aspect angle or orientation

with respect to the enemy tracking radars line of sight (LOS): ratio of radar

wavelength to target size: polarisation of transmit and receive antennas of the enemy

radar: surface quality: and constitution of the target. Methods required to control RCS

depend critically on the size of the electrically conducting component being

illuminated compared with the wavelength of the radar signal illuminating it.

If the wavelength of the signal is much less than the physical size of the component,

and if it is smooth enough, radar waves reflect much as a mirror reflects light.



216

Table B.1: - Typical radar threats.

Radar system. Frequency (GHZ). Wavelength (cm).

Early Warning. 0.15 - 0.2 150.0 – 200.0

3.0 - 4.0 7.5 – 10.0

Ground control

interception (GCI) 2.0 – 3.0 7.5 – 15.0

Height finders 2.0 – 7.0 4 – 15.0

Aircraft 8.0 – 20.0 1.5 – 4.0

Air – to – air missiles 10.0 – 20.0 1.5 – 3.0

SAM transportable.

Acquisition 0.15 – 3.0 10.0 – 200.0

Tracking 5.0 – 10.0 3.0 – 6.0

SAM mobile.

Acquisition 2.0 – 6.0 5.0 – 16.0

Tracking 5.0 – 13.0 2.3 – 6.0

Radar – guided anti

aircraft artillery AAA 14.0 – 16.0 1.8 – 2.0

(Reference 25: - page 137, Fundamentals of Fighter Design: by Whitford Ray:

Published by Airlife Publishing Limited UK 2000).

The RCS of an aircraft is determined by the magnitudes of two distinctly different

contributions: -



217

(a) The aircrafts shape (Outer Mould Lines – OML) both overall and in detail

(including surface quality). (Realistic reductions in combat aircraft size have an

insignificant effect on their RCS): (b) The electromagnetic properties of the

airframe materials.

The major contribution features to RCS for conventional non-stealthy aircraft are as

follows: - engine compressor faces (forward) and turbines (aft) in radar LOS due to

their Doppler signature: engine air intakes and diverter plates: external stores,

including missiles seeker heads: wing leading edges: corner reflections at

intersections of fins and tailplanes: wing planform viewed from above or below:

radome and bulkhead, if transparent to illuminating radar: cockpit, including cavity

effect due to the very large number of corner reflectors: engine nozzle if viewed from

rear: flat slab-sided fuselage when viewed from side elevation. All of these features

are present on the BAE Systems Tornado GR-1A figure B1 below; Boeing F-15E;

and General Dynamics (now Lockheed Martin) F-111‟s rendering them holey

inadequate, except against the most primitive of enemy defences. (Ref 25)

Figure B.1: -Tornado GR-1A, externally similar to the GR-4 is the only RAF

deep strike aircraft and the lack of any LO capabilities contributed to aircraft

losses in the first Gulf Wars. Source: - BAE Systems.

The same features are equally evident on the F-15E Strike Eagle figure B.2, the

USAF‟s current strike asset, and the RAAF F-111C shown in figure B3.

Active sensors and

ECM systems

detectable emissions

External weapons, and

fuel high RCS signature

Afterburning engines high IR

signature and turbine face visible

Right angle sides and has

no planform alignment

Engine compressor

face visible to radar Tall fin at right angle to

horizontal tails



218

Figure B.2: -F-15E Strike Eagle the USAF equivalent to the RAF Tornado

developed from the F-15D air superiority fighter. Source: - USAF Academy.

Figure B.3: -F-111E the USAF equivalent to the RAAF F-111C from which it was

developed, note the F-111‟s were retired from USAF service in the mid 1990‟s.

Source: - USAF Academy.

Smaller contributions to RCS for a non stealthy aircraft are as follows: - fuselage in

head on view: wing leading edge and control surface gaps which cause scattering: local

air inlets e.g. for cooling and air conditioning: local surface protuberances, even the

Two seat non coated

canopy

External stores only under wing

and under fuselage

No blending and no planform alignment

Twin fins at right angle to

horizontal tails

No blending and no planform alignment

Two seat non coated

canopy

External stores, under wings



219

smallest protuberances cannot be ignored and each may become resonant at a different

frequency: long thin fairings including missiles fins and tailplanes.

Note if the major contributors to RCS are carefully dealt with then the previously minor

ones become important, therefore stealth is difficult to achieve but easy to lose through

lack of attention to detail. Aircraft shaping is useful over a wide range of radar

frequencies but over a limited range of aspect angles. Typically, for fighter aircraft, the

forward cone of angles is of greatest interest and, hence large returns can be shifted out

of this sector into the broadside directions. Aircraft can be shaped to ensure that most

radar waves will be scattered and not reflected back to the transmitter. Leading and

trailing edges of the wings, control surfaces, inlet lips, door gaps, etc., can be aligned to

ensure that the energy that is unavoidably, reflected back to the transmitter is

concentrated into a few spikes, as shown in figures B4 and B5 below.

This will give the enemy radar one good return when the alignment is ideal, but a

much weaker return on subsequent scans. (Ref 25)

Figure B.4:- Spike alignment or planform alignment on the F-35C, the alignment

of the leading and trailing edges of the wings, control surfaces, inlet lips, door

gaps, etc., are aligned to ensure that the energy that is reflected back to the

receiver is concentrated to 8 spikes away from the forward cone, also the canopy

is coated with a tin iridium oxide to make it opaque to radar, and the radome is a

selective wave pass design.

Spike

Alignment

Spike

Alignment



220

Figure B.5:- The same spike alignment or planform alignment is applied to the

fuselage sides and vertical tails on the F-35C, which have the same common

alignment angle of 270 to ensure that the energy that is reflected back to the

receiver is concentrated away from the illumination source.

Two different approaches to aircraft shaping to reduce RCS have been developed

which primarily resulted from the amount of computing power available for analysis

when they were formulated: -

The first approach was the faceted configuration which used flat panels to minimise

normal reflections back to the illuminating radar. The major work on faceting for RCS

reduction was undertaken at Lockheed Skunk Works by Denys Overholser who

applied a work on diffraction published in 1962 by the Russian physicist Ufimtsev, to

the problem of RCS reduction. This lead to the development of a computational

method of predicting the RCS of two dimensional shapes built up from a series of flat

surfaces. This eventually gave rise to the Have Blue and F-117A Nighthawk aircraft

shown in figure B.6 below. (A detailed description of this approach is given in

appendix b of Reference 27).

The second approach was the use of a compact, smoothly blended external geometry

to achieve a continuously varying curvature, as employed in the Northrop Grumman

B-2 and the Lockheed Martin F/A-22A and the F-35 Joint Strike Fighter, this required

far grater predictive ability and enormously increased computational capacity. (A

detailed description of this approach is given in the appendix of Reference 26).

Spike

Alignment



221

Figure B.6:- F-117A Nighthawk illustrating facetted construction to reduce RCS

and slotted engine exhaust to reduce IR signature as well as rear quarter RCS.

By the late 1950‟s engineers in the field of radar signature reduction realised that the

very large RCS reductions necessary to achieve any operational benefit could not be

accomplished simply by coating an otherwise conventional aircraft with radar

absorbent materials (RAM) (a lesson seemingly forgotten outside the USA for

example in the Dassault Rafale), Many physically small and apparently insignificant

features of an aircraft generate radar returns that are still quite detectable. Indeed a

very low RCS must be designed into an aircraft from the outset, with rigorous

attention paid to all three elements of RCS control: overall shaping of the airframe and

its components, special detail treatments, and RAM. (Ref 13)

The Advanced Interdiction Aircraft being derived from the F-35 JSF family of aircraft

will capitalise on the experience gained within Lockheed Martin and Northrop

Grumman, from development and production of their legacy stealth aircraft namely

the: - Lockheed Martin F-117A Nighthawk: F/A-22A Raptor: and the Northrop

Grumman B-2A Spirit. Additional experience from in service evaluation of these and

other low-observable UCAV projects within the F-35 team have been utilised to make

the F-35 more affordable and more easily supportable in front line units. The

Advanced Interdiction Aircraft will employ the same low observable technologies as

the F-35 family demonstrated on the high – fidelity SigMA model shown in figure B7.

A critical objective of the F-35 and the AIA program is to produce a stealthy aircraft

that stays stealthy in severe combat conditions. This goal is best achieved by building

an aircraft that is hard to damage requiring durable battle damage tolerant structure

and making sure that anything except the most severe damage will not significantly

degrade the RCS. The RCS of the FB-24 AIA would be validated by a similar test

programme to that detailed above for the F-35.

Slotted exhaust nozzle

Grill covered intakes

High drag facetted OML



222

Figure B.7: - The SigMA model about to undergo RCS tests at the Helendale,

California facility. Source: - Lockheed Martin Press Release Office.

The design must also ensure that any required repairs are rapid and can be performed

at front line units on the flight line. The SigMA testing program demonstrated this and

the AIA will build upon the experience, as part of the F-35 family.

The SigMA models shown in figures B7 and B8 were full scale aircraft models

complete with a representative internal structure, and high – fidelity features including

removable doors and access panels, canopy transparency, cockpit interior, external

lights, air data probes, engine components, edges, re-positional control surfaces,

antenna apertures, radar array, and a flight capable nose – cone. Testing began in

February 2000 and included measuring aircraft RCS and the performance of various

antennas on the SigMA model aircraft. Tests also demonstrated the robustness of

supportable low – observable (LO) materials and their repair.

After baseline testing, several doors and panels were intentionally damaged and later

repaired, then installed on the model and RCS measurements were taken to determine

the impact of the defects and the effectiveness of the repairs.

The inflicted damage, more than three dozen significant defects represented in types

and frequency the cumulative effect of more than 600 flight hours of military aircraft

operations.

When engineers overlaid the RCS curves of the undamaged configuration on the

damaged and repaired configurations they found it extremely difficult to detect any

changes between the two configurations. These repairs were conducted quickly and

easily with engineers repairing all of the damage within a single eight hour shift. A

similar repair on a legacy B-2A would be estimated to require more than 72 hours.

RCS testing was completed by early October 2000, validating the F-35 OML, the

resilience of the LO materials, and previewing its cost – savings potential.



223

Figure B.8: - SIGMA Model on test pole. Source: - Lockheed Martin Press

Release Office.

In further efforts to reduce maintenance and supportability costs and maintain

signature continuity in October 2000, Lockheed Martin demonstrated a new

generation of “paint – less” aircraft covering film on a fully covered F-16 which made

a series of successful test flights including one at Mach 1.8 which is the highest speed

ever reached in an appliqué flight test. Aircraft appliqués consist of paint –

replacement adhesive films designed to bring savings in production costs, support

requirements, disposal costs and importantly savings in aircraft weight. (They also

offer significant environmental advantages since military painting operations are a

significant source of hazardous – material emissions).

(Reference 6:- Lockheed Martin F-35 Joint Strike Fighter: The Universal Fighter: by

Harkins H: Published by Centurion Publishing UK 2004)

The Advanced Interdiction Aircraft will use this appliqué coating technology to enable

frequency tuned material application at reduced weight, and ensure signature

continuity over the airframe.

2.3.4 Infra-red Signature Control.

Passive (non-emitting) IR detection targeting devices rely on contrasts between hot

parts on an airframe such as jet pipes and surfaces subjected to kinetic heating, and the

background radiation. Many IR devices operate in the 8 to 13 micron band because

this is the most IR transparent band in the atmosphere. In engine exhaust, carbon

dioxide produces most of the IR signature at 4.2 microns, so most modern IR sensors

can “see” at two different wavelengths (medium:- 3 to 5 microns and long:- 8 to 14

microns) to provide good target discrimination.



224

The radiated IR energy is proportional to the fourth power of absolute temperature (i.e.

T4). With engine turbine entry temperatures currently in the order of 1,627

o C, and

rising, the rear fuselage of a fighter is the greatest source of IR radiation. With

afterburner on, IR emissions are vastly increased. Moderate stagnation temperatures

(which vary with Mach number squared) are inevitable due to kinetic heating on

leading edges of fighters flying at high Mach numbers.

As the stealthiness of fighters improves, the exhaust plume of the missiles launched by

the fighters becomes a contributor to IR detection. Low – visibility plumes will

minimise detection of both launch platform and missile (vitally important for the

ALOSNW and the AIA). There are several ways of reducing the IR signature of a

fighter aircraft which are as follows:-

The ability to supercruise (cruise at supersonic speeds without afterburner)

limited to aircraft with a thrust to weight ratio of 1:1 or above like the F/A-

22 Raptor and YF-23, this reduces the temperature of the nozzles: moreover,

supercruising allows the pilot to engage on his terms, increases weapons

envelopes, minimises exposure to SAM threats and not only stretches

combat radius but forces an adversary to expend his own fuel in order to get

his aircraft to an energy state where he can engage it.

Although not currently available to the F-35 family, the JSF 119-614 engine

has demonstrated installed thrust of 50,000lbs and has further growth

potential, and also variable cycle engines like derivatives of the GE YF-120

ATF engine could offer the potential of thrust to weight ratios of 1:1 in the

future. The Advanced Interdiction Aircraft would also benefit from such

developments and an engine based on the GE VCE technology should be

explored for supercruise capability.

Use of a high bypass-ratio (BPR) engine to mix in cold air to reduce exhaust

temperature, however a BPR of more than 0.4 conflicts with the requirement

of high dry thrust to achieve supercruise.

Use of a curved jet pipe to mask the hot turbine stages.

Use of two – dimensional nozzles (which increase the surface area of the

exhaust plume) or ejector nozzles (which mix in ambient air) to increase the

rate of cooling.

Increase cooling of the outer skin of the engine bay or insulate to reduce the

temperature of the engine bay outer skin.

Use a curved air intake (the F-35 and AIA use a bifurcated duct) to mask

forward emissions from the engine.

Limit maximum supersonic speed to reduce IR signature due to kinetic

heating (one beneficial side effect of reducing the AIA dash speed).



225

Figure B.9: - The Pratt & Whitney LON nozzle for use on the FB-24, undergoing

ground test during the Concept Demonstration Phase of the JSF program.

Source: - Pratt & Whitney Press Release Office.

Figure B.10: - The General Electric AVEN nozzle design of the type to be

employed on the A-24, not shown in this picture are the chevron trailing edge

tips the nozzle segments which reduce rear illumination radar returns

(Reference 15).



226

The FB-24 in common with the F-35 family will employ the Pratt & Whitney Low

Observable convergent / divergent exhaust nozzle on the F-135 engine (ground tested

as shown in figure B9 below), or the GE Low Observable Axisymmetric Nozzle

(LOAN) on the F-136 engine (ground tested on a Lockheed Martin F-16C for over

500-hours) (References 15 and 16) as will the A-24, although employment of the GE

Aircraft Engines AVEN axisymmetric vectoring nozzle with exhaust vectoring in any

direction up to 20 degrees from the centre axis shown in figure B10 was also

considered.

The two different nozzle types considered for the FB-24 and the A-24 would be

tailored to each airframe rear fuselage design to optimise the performance and

survivability of each aircraft, but these nozzles would be common from the turbine aft

face up to the throat (the point of minimum cross-sectional area). Although the Pratt

& Whitney nozzle would not have a thrust vectoring capability the GE LOAN /

AVEN would have and the A-24 would require such a nozzle due to its direct

penetration role, and greater need for threat avoidance, the FB-24 would also benefit

from increased manoeuvrability.

Figure B.11: - The 2-D thrust vectoring nozzles of the F/A-22A shown here to

good effect, the top and bottom chevron tipped flaps move up and down in

unison to direct the engine thrust imparting high agility to this large aircraft

(Source: - Lockheed Martin Press Office).

The 2-D vectoring nozzle system as employed on the F/A-22A shown in figure B11

above was considered too heavy and complex and inappropriate for the missions

envisioned for the FB-24 and A-24, which would evade rather than engage the enemy

and therefore would not require the agility of an air dominance fighter.



227

Appendices C: - Details of the F-35 Diverterless Supersonic Intake.

A fighter inlet must provide the engine or engines with high-quality airflow over a

wide range of speeds, altitudes, and manoeuvring conditions while accommodating the

full range of engine airflow from idle to maximum military or afterburning power.

The inlet designer must also consider the constraints imposed by configuration

features, such as nose landing gears, weapon bays, equipment access panels, and

forebody shaping. The designer must produce the lowest drag, lowest weight, lowest

cost, and highest propulsion performance solution. It must also meet stringent low

observable requirements.

Historically, inlet complexity has been a function of the top speed for a fighter aircraft.

Higher Mach numbers require more sophisticated devices for compressing supersonic

airflow to slow it down to subsonic levels before it reaches the face of the engine. (Jet

engines are not designed to handle the shock waves associated with supersonic

airflow.) These compression schemes involve the conversion of the kinetic energy of

the supersonic air stream into total pressure on the compressor face of the engine.

Speeds over Mach 2 generally require more elaborate compression schemes like those

of the F-14, F-15, and MiG-25 Foxbat MiG-31 Foxhound.

The F-15 inlet, for example, contains a series of movable ramps and doors controlled

by software and elaborate mechanical systems. The ramps move to adjust the external

and internal shape of the inlet to provide optimum airflow to the engine at various

aircraft speeds and angles of attack.

Doors and ducting allow excess airflow to bypass the inlet. The entire system results

in high radar returns and a high weight penalty for the aircraft, signature reduction was

not a priority when the “teen series” of aircraft was designed and built.

Fighter inlet designers must also account for boundary layer air which is a layer of low

- energy air that forms on the surface of the fuselage at subsonic and supersonic

speeds. (These layers also form on the inlet compression surface). This layer of

relatively slow moving, turbulent air, can create havoc when disturbed by the shock

waves created by the inlet. This results in unwanted airflow distortions at the engine

face. If the shock wave / boundary layer interaction is severe enough, the engine will

stall.

An important consideration for the Advanced Interdiction Aircraft is that this

boundary layer increases in thickness with speed and forebody length, as the AIA

seeks a 3ft increase in forebody length for the second crew member, later CFD

analysis should be directed to investigate the impact on DSI size and effectiveness, the

intake position may need to change as on the F/A-22 or the original X-35 / 230-2

OML intake configuration may have to be adopted with its inherent weight penalty (to

be addressed in future Trade Studies).



228

Fighter designers‟ deal with this boundary layer phenomenon by redirecting the layer

before it reaches the engine by placing the inlet away from the boundary layer in the

free stream, where airflow is unaffected by the boundary layer phenomenon. On the F-

16, a structure called a diverter provides a 3.3-inch gap between the fuselage and the

upper lip of the intake. The size of the gap is directly proportional to the thickness of

the boundary layer at maximum speed of the F-16. Other fighters remove boundary

layer airflow with combinations of splitter plates and bleed systems. The latter

redirecting the unwanted airflow through small holes in the compression ramps to

bleed ducts within the inlet. All of these systems result in radar waves being either

returned from the splitter plate / forebody cavity or from the compression ramps.

The Diverterless Supersonic Inlet bump shown in figure C1 functions as a

compression surface and creates a pressure distribution that prevents the majority of

the boundary layer air from entering the inlet at speeds up to Mach 2 as described in

the Aerodynamic V3 slide shown in figure C2.

In essence, the DSI dispenses with complex and heavy mechanical systems as well as

significantly reducing the RCS of the aircraft.

The DSI concept was introduced to the F-35 (then JAST / JSF) program as a trade

study item in 1994, and was compared with the traditional conical translating inlet

seen on F-111‟s. After CFD analysis supported by wind tunnel testing full-scale flight

testing was conducted using an F-16 at Mach 2, full scale testing culminated in the X-

35 Concept Demonstration Phase with both CDA X-35 aircraft built with the DSI

configuration and demonstrating its capabilities throughout the complete test program.

Figure C.1: - F-35 SDD mock up aircraft showing the DSI and a cut back intake

lip compared with the X-35 aircraft as a weight reduction measure. Source: -

Code One magazine reference 17.



229

WaveriderWaverider--like like ““BumpBump”” diverts diverts

boundary layer using pressure boundary layer using pressure

gradientgradient

CFD tool advances allowed for CFD tool advances allowed for

integration into todayintegration into today‟‟s vehicless vehicles

STEPS:STEPS:

Define 3Define 3--D Compression Surface D Compression Surface

From From ““Virtual ConeVirtual Cone”” CFD SolutionCFD Solution

•• Early:Early: Traditional ConeTraditional Cone

•• SOA:SOA: Isentropic Cone at Isentropic Cone at

Develop Develop ““CenterlineCenterline”” GeometryGeometry

•• Compression Surface / Compression Surface /

Shoulder / Diffuser Fairing Shoulder / Diffuser Fairing

IntegrationIntegration

•• Cowl / ACowl / Aii / A/ A

tt IntegrationIntegration

Integrate Complete Inlet / Integrate Complete Inlet /

Forebody Forebody

•• Forebody / Aperture / Duct Forebody / Aperture / Duct

IntegrationIntegration

•• Real Aircraft ConstraintsReal Aircraft Constraints

* Protected by U.S. Patents 5,749,542 & 5,779,189* Protected by U.S. Patents 5,749,542 & 5,779,189

ADVANCED INLET INTEGRATION

Diverterless Inlet Technology*

Conic

al S

hock

Conic

al S

hock

Stream Lines

Stream Lines

CompressionSurface

CompressionSurface

FLOWFLOW

Virtual ConeVirtual Cone

1

ConeCone

Comp Surface

Comp Surface

TransitionShoulder

TransitionShoulder

DiffuserFairing

DiffuserFairing

CowlCowl

2 3

NACA RM E56L19 (1957)NACA RM E56L19 (1957)

Compression OnlyCompression Only Compression + BL DiversionCompression + BL Diversion

Figure C.2: - The operational principals of the DSI and CFD representation of

the boundary layer flow fields around DSI with the original X-35 intake lip

configuration. Source AIAA AeroDYNAMIC V3.

A more detailed description of the flight test DSI program on the F-16 can be found in

reference 17 which was the source of the material presented in this section.

(Reference17:- Pages 9 to 13, Code One magazine: Volume15 number 3: by Hehs

Eric: Published by Lockheed Martin Aeronautics Company 200)

Appendices D: - Current F-35 family of aircraft.

The current F-35 Joint Strike Fighter family consist of three aircraft variants for three

United States military air arms and the United Kingdom armed forces, which have

common basic configurations and use cousin parts in most cases, are outlined below

in terms of lead dimensions: performance: and weight and are shown in figures D1 -

D6.



230

Figure D.1: - F-35A USAF (CTOL Variant). Source: - Reference 3:- Document

number 17521 Briefing by Brig Gen J. Hudson.

F-35A (CTOL Variant): - This is the USAF variant with 9g manoeuvring capability

and replaces the F-16 and A-10 (current orders are for 1763 aircraft). Dimensions

are: - Wing span 35 ft: length 51.1 ft: and wing area 460 ft2. Empty weight is: -

27,395 lbs. Internal fuel capacity is: - 18,498 lbs. Combat radius is: - greater than

590 nautical miles.

Figure D.2: - X-35A CDA for F-35A in fight 2001:-Source Authors Private

collection.



231

Figure D.3: - F-35B USMC / UK RAF and RN (STOVL Variant). Source: -

Reference 3:- Document number 17521 Briefing by Brig Gen J. Hudson.

F-35B (STOVL Variant): - This is the USMC / U.K. RN & RAF variant to replace

the Harrier AV-8B / Mk - 9 and the F/A - 18 Hornet C/D aircraft (current orders are

for 759 aircraft). Dimensions are: - Wing span 35 ft: length 51.1 ft: and wing area

460 ft2. Empty weight is: - 30,697 lbs. Internal fuel capacity is: - 13,326 lbs.

Combat radius is: - greater than 450 nautical miles.

Figure D.4: - X-35B CDA for F-35B Edwards Air Force base tests 2001:-Source

Authors Private collection.



232

Figure D.5: - F-35C USN (CV Variant). Source: - Reference 3:- Document

number 17521 Briefing by Brig Gen J. Hudson.

F-35C (CV Variant): - This is the US Navy aircraft carrier variant which will

replace the F/A - 18 Hornet C/D and the F - 14D Tomcat (current orders are for 480

aircraft). Dimensions are: - Wing span 43 ft: Length 51.4 ft: Wing area 620 ft2.

Empty weight is: - 30,618 lbs. Internal fuel capacity is: - 19,100 lbs. Combat

radius is: - greater than 600 nautical miles. Maximum payload all internal and

external stores stations used: - 21,400lbs.

Figure D.6: - F-35C in maximum stores configuration: - Source Reference 3:-

Document number 17521 Briefing by Brig Gen J. Hudson.



233

Appendices E: - Design Supplement.

The objective of this appendices was to cover aspects of the FB-24 and A-24 design

which could not be covered in the bulk of the report and the following aspects of the

design are covered herein: -

1) Structural design amendments and concept completion with weight analysis:

2) Supersonic range and endurance and mission stage fuel burn for NB2:

3) Aerodynamic and performance methods of the AeroDYNAMIC toolset:

4) Comparison of as designed aircraft with the RFP targets.

E-1.1:- Structural design amendments and concept completion

The author was not satisfied with the initial proposal for a large MOSC ingress /

egress hatch of 2.0m wide and 1.2m in length, he felt that this large cut out would

weaken the front fuselage and would be difficult to seal effectively and ample

provision for MOSC emergency ejection could be provided by a smaller aperture

which would land on additional shoulder longerons framing the MOSC compartment.

This new hatch would have the advantage of requiring a smaller and lighter actuation

system, a lighter structure, and less risk of impacting the airframe on release which

could result in debris impacting the aircrew during emergency decent after ejection.

This new hatch is shown in figure E-1 below and was 0.88m wide by 1.6m in length,

with all round pressurization sealing. The hatch was hinged at the rear and rotated

around a single hinge bolt designed to fail under hatch emergency release loads i.e.

when the hatch was opened into the airflow over the forward fuselage over the canopy

the bolt would shear and the hatch would separate from the aircraft. The ejection

sequence would be MOSC first followed by the pilot. The hatch would be secured at

the front face by two explosive release automatic locking bolts, of the same type used

to attach the actuator to the hatch, and the front face would seal on the pilots canopy

frame. The integration of this hatch with the forward fuselage is shown in figure E-2

below were the hatch is shown open to 450 in normal operation this would open to 80

0

to enable ingress and egress of the MOSC in full NBC fight clothing, as well as

conventional Night Vision Goggles and the Advanced Virtual Environment Helmet,

proposed for the FB-24. The original proposal to chevron the forward face of the

hatch in the same manner as the F-117A canopy frame was dropped because this face

would be completely obscured from radar illumination by the pilots canopy obviating

the need for treatment beyond a RAM loaded „p – seal‟. This concluded the forward

fuselage design proposal for the FB-24 airframe. The A-24 requirements are covered

in figures E-3 and E-4 below.

The second major design issue reflected in the common fuselage structural design of

both the FB-24 and A-24 airframes was the load paths in the rear fuselage in which all

of the boom loads were carried by the single fuselage closure frame. This was clearly

unsatisfactory as the resulting frame would be massive and very heavy so to resolve

this, the author proposed a kicked inboard keel boom attaching to the shoulder

longeron and lower longeron of the aft fuselage, as well as a continuous outboard keel

as shown in figures E-5 and E-6 below.



234

Figure E-1: - Proposed MOSC hatch as described above made from RTM

substructure and fibre placement skin all in Carbon PMR-15 high temperature

composite.

Figure E-2: - Proposed MOSC hatch integration with forward fuselage as

described above with frame stations (FS) numbered in metres.

0.88m

1.6m

Co-Bonded Stub frames Hatch sealing and

landing longerons

Hatch attachment longerons

Hatch sealing and

landing longerons

Frame back-up stiffeners and

hatch attachment fixtures.

Hatch open at 450

FS2.51

FS2.92

FS3.35

FS3.71

FS4.09

FS4.46

FS4.88 FS5.35

FS5.70 FS6.12 FS6.43



235

AI

Mo

du

le

Op

tical

tra

ckin

g

syste

m

180

0 f

ield

of

reg

ard

po

rt /

sta

rbo

ard

an

d -

16

0 t

o 1

20

0 in

ve

rtic

al p

lan

e.

Ele

ctr

o

Op

tical

Ta

rgeti

ng

S

yste

m

EO

TS

mo

un

ted

beh

ind

a f

ac

ete

d s

ap

ph

ire

win

do

w.

AI M

od

ule

wit

h E

CS

inco

rpo

rate

d

Sele

cti

ve

RF

fre

qu

en

cy a

nd

wave l

en

gth

LO

rad

om

e.

Ch

evro

n A

ES

A r

ad

ar

mo

un

t fr

am

e

LO

co

ate

d f

ield

of

reg

ard

240

0

Sta

rbo

ard

Can

op

y

su

pp

ort

lo

ng

ero

n

Po

rt F

orw

ard

ch

ine lo

ng

ero

n

Po

rt C

an

op

y

su

pp

ort

lo

ng

ero

n

Po

rt In

tak

e c

hin

e

lon

gero

n „C

‟ secti

on

Po

rt F

orw

ard

fu

sela

ge

sh

ou

lder

lon

ge

ron

Carb

on

P

MR

-15 all fr

am

es w

ith

inte

gra

l sti

ffen

ers

: -

F1

fu

el

tan

k

bo

un

dary

fra

me

s:

mati

ng

fra

me:

inte

rface

co

nn

ecto

r su

pp

ort

fram

e.

Fig

ure

E-3

:- A

-24 F

orw

ard

fu

sela

ge

stru

ctu

ral

lay

ou

t an

d p

osi

tion

ing o

f A

I m

od

ule

.



236

AI Track and scan module.

The down selected configuration for the modular AI optical tracking system was

based on the requirement to give the operator a pilot eye view with a full 1800 field of

regard and a 160 over nose view angle, imparting greater awareness on landing, take-

off, and in fight refuelling shown below for EFA in figure E-4, of the A-24 UCAV.

This OTS field of regard combined with the chin mounted Electro Optical Targeting

System (EOTS), all-digital infra-red system, offers a much grater passive situational

and target awareness than the Predator for example which has been compared to

flying through a drinking straw. The OTS is shrouded by a multi layer tin iridium

oxide coated opaque to radar optically perfect bowless canopy.

Figure E-4: - A Spanish Eurofighter Typhoon refuelling from a C-130 tanker

using the hose and drogue technique advocated for both FB-24 and A-24

aircraft.

Although autonomous air to air refuelling should be fairly simple with the boom

method in which the recipient aircraft is passive station keeping role, and the boom is

flown from the tanker into the receptacle, it limits the number of aircraft which can be

refuelled to one per tanker. The hose and drogue method is much more demanding

with the recipient aircraft in the active role of achieving hook on and station keeping,

but permits up to three aircraft to be refuelled at the same time (all be it with a

reduced fuel transfer rate) therefore a large field of regard is essential for this

operation to be undertaken by an off-board pilot. The AI interface with the aircraft

systems would be through a common (manned / unmanned) optical FBL data bus

connection housed in the lower forward fuselage, enabling plug and play

interchangeability of AI units.



237

Fig

ure

E-5

:- F

B-2

4 /

A-2

4 C

om

mon

sim

pli

fied

aft

fu

sela

ge

stru

ctu

ral

layou

t to

p v

iew

.

Sta

rbo

ard

T

i fu

ll

len

gth

co

nti

nu

ou

s

„T‟

sec

tio

n

sh

ou

lder

lon

ge

ron

s 7

0m

m w

ide f

lan

ge b

y 5

0m

m d

eep

Sta

rbo

ard

Ti

full l

en

gth

co

nti

nu

ou

s „

I‟ s

ecti

on

ke

el

beam

70m

m w

ide f

lan

ge f

ull d

ep

th w

ith

sp

igo

t b

eari

ng

ho

usin

g

Po

rt C

arb

on

BM

I N

ac

elle

skin

P

ort

T

i fu

ll

len

gth

co

nti

nu

ou

s

„C‟

se

cti

on

low

er

lon

gero

ns 3

0m

m w

ide b

y 1

00m

m d

eep

Po

rt T

i fu

ll l

en

gth

aft

fu

sela

ge „

C‟

se

cti

on

ch

ine lo

ng

ero

n

Po

rt T

i „I‟

se

cti

on

kic

ked

keel

be

am

70m

m w

ide f

lan

ge f

ull d

ep

th

Po

rt T

i in

bo

ard

„C

‟ s

ecti

on

ch

ine lo

ng

ero

n

Ru

dd

erv

ato

r sp

igo

t h

ou

sin

g t

ub

es.

Aft

fu

se f

ull f

ram

es

F-1

36

-2 E

ng

ine

UP

AF

T



238

Fig

ure

E-6

:- F

B-2

4 /

A-2

4 C

om

mon

sim

pli

fied

aft

fu

sela

ge

stru

ctu

ral

layou

t u

nd

ersi

de

vie

w.

Aft

fu

se

bo

w

fram

es

a

s

a

rem

ovab

le c

en

tre s

ecti

on

of

the

main

fr

am

e

for

en

gin

e

extr

acti

on

FR

-5 F

uel ta

nk b

ou

nd

ary

fra

mes

Sta

rbo

ard

bo

om

actu

ato

r b

ays

F-1

36

-2 E

ng

ine

FL

-5 F

uel ta

nk b

ou

nd

ary

fra

me

s

Po

rt b

oo

m a

ctu

ato

r b

ay

s

NB

:- A

ft f

usela

ge r

ed

esig

n p

rovid

es

co

nti

nu

ou

s

load

p

ath

s

inte

rfa

cin

g

wit

h t

he c

en

tre f

us

ela

ge

keels

.

FS

14.9

2

FS

15.4

1

FS

14.3

9

FS

15.8

9

FS

16.3

7 F

S17.0

8 R

F

S17.9

4 R

FS

17.0

8 L

FS

17.9

4 L

A

FT

UP



239

Weap

on

s b

ay l

oad

path

keel

En

gin

e a

ttach

men

t fr

am

es A

l fo

rward

of

max t

em

pera

ture

zo

ne a

nd

Ti at

tru

nn

ion

att

ach

men

t / m

ate

fra

me

Un

derc

arr

iag

e lo

ad

carr

yin

g f

ram

es

En

gin

e t

run

nio

n

att

ach

men

ts

En

gin

e lu

g a

tta

ch

men

t

Ti

sid

e w

alled

weap

on

s b

ays a

ttach

ed

to t

he f

ram

es f

orm

ad

dit

ion

al lo

ad

path

Al

full l

en

gth

co

nti

nu

ou

s „

C‟

se

cti

on

lon

gero

ns 7

0m

m b

y 5

0m

m

So

ft b

uild

jo

int

NB

:-

Cen

tre

fus

ela

ge

o

nly

m

ajo

r

ch

an

ge

was

ad

dit

ion

o

f w

eap

on

s

bay

inb

oard

w

all

load

p

ath

keel

wh

ich

acts

as

a

fusela

ge

ben

din

g

load

dis

trib

uto

r o

utb

oard

of

the m

ain

keel

be

am

s an

d m

ate

s w

ith

th

e aft

fusela

ge in

bo

ard

bo

om

lo

wer

ke

els

.

Fig

ure

E-7

:- D

etail

ed s

tru

ctu

ral

layou

t u

pd

ate

of

the

com

mon

cen

tre

fuse

lage

of

the

FB

-24 /

A-2

4. T

op

vie

w s

how

n.

Lan

d

based

a

rreste

r

ho

ok a

ttach

men

t fr

am

e

FS

8.3

8

FS

9.5

2 F

S10.6

2

FS

11.6

0

FS

7.1

2

FS

12.5

8

FS

13.1

5

FS

13.9

7



240

Fig

ure

E-8

:- D

etail

ed s

tru

ctu

ral

layou

t m

od

el o

f re

vis

ed F

B-2

4 c

on

figu

rati

on

.

FW

D f

usela

ge b

uild

mo

du

le

RT

M c

arb

on

PM

R-1

5

Weap

on

s b

ay

s T

i S

PF

/DB

Cera

mic

/ S

tru

ctu

ral

RA

M le

ad

ing

ed

ge

fla

p

Cera

mic

/ S

tru

ctu

ral

RA

M f

lap

ero

n

Aft

fu

sela

ge b

uild

mo

du

le in

Ti

Ti / S

tru

ctu

ral R

AM

rud

derv

ato

rs

Carb

on

BM

I w

ing

sk

in t

hic

kn

es

s

ran

ge 2

0m

m a

t ro

ot

– 4

mm

at

tip

W

ing

str

uctu

re T

i

Cen

tre f

use

lag

e b

uil

d

mo

du

le in

Al

an

d T

i C

arb

on

BM

I fu

sela

ge

skin

th

ickn

ess

5.0

8m

m

Ti

fire

sh

ield

n

ot

sh

ow

n

FS

10

149 –

to

–F

S15

893.



241

Fig

ure

E-9

:- D

etail

ed s

tru

ctu

ral

layou

t m

od

el o

f re

vis

ed A

-24 c

on

figu

rati

on

wit

h 7

5%

com

mo

nali

ty w

ith

FB

-24.

FW

D f

usela

ge b

uild

mo

du

le

RT

M c

arb

on

PM

R-1

5

Weap

on

s b

ay

s T

i S

PF

/DB

Cera

mic

/ S

tru

ctu

ral

RA

M le

ad

ing

ed

ge

fla

p

Win

g s

tru

ctu

re T

i

Cera

mic

/ S

tru

ctu

ral

RA

M le

ad

ing

ed

ge

fla

p

Carb

on

BM

I w

ing

sk

in t

hic

kn

es

s

ran

ge 2

0m

m a

t ro

ot

– 4

mm

at

tip

Cen

tre f

use

lag

e b

uil

d

mo

du

le in

Al

an

d T

i

Cera

mic

/ S

tru

ctu

ral

RA

M f

lap

ero

n

Aft

fu

sela

ge b

uild

mo

du

le in

Ti

Ti / S

tru

ctu

ral R

AM

rud

derv

ato

rs

Carb

on

BM

I fu

sela

ge

skin

th

ickn

ess

5.0

8m

m

Ti

fire

sh

ield

n

ot

sh

ow

n

FS

10

149 –

to

–F

S15

893.



242

Fig

ure

E-1

0:-

Det

ail

ed s

tru

ctu

ral

layou

t m

od

el o

f re

vis

ed c

om

mon

FB

-24 /

A-2

4 w

ing.

Rib

1A

Rib

1B

Rib

1C

Rib

1D

Rib

1E

Rib

2A

Rib

2B

Rib

2C

Rib

2D

Rib

3A

Rib

3B

Rib

R (

A)

Rib

R (

B)

Rib

R (

C)

Rib

R (

D)

Rib

R (

E)

Rib

R (

F)

Tip

Rib

Au

x R

ib 1

Au

x R

ib 2

(a)

/ (b

)

Au

x R

ib 3

Fw

d S

par

Sp

ar

1

Sp

ar

2

Sp

ar

3

Sp

ar

4

Sp

ar

4

Sp

ar

5

Sp

ar

5

Aft

Sp

ar

Stu

b S

pa

r



243

Figure E-11: - Structural joint philosophy for rib to spar metallic joints.

Figure E-12: - Proposed structural joint philosophy for Ti spar to carbon PMR-

15 rib joints to be employed in the carbon composite wing torsion box

substructure trade study illustrated below in figure E-13.



244

Figure E-13:- Proposed weight reduction trade study to using carbon PRM-15

for wing torsion box substructure using the layout defined in figure E-10.

Figure E-14:- The A-24 estimated total fuel capacity 25,135lbs equal to 16%

greater fuel capacity of the two manned FB-24, and estimated GTOW of

48,200lbs.

Ti wing boundary and carbon PMR-15 sub-structure with multi spar layout to resist buckling of

skins with long thin panels.



245

E-1.2:- Structural design weight estimation

The structural weight measurement for each major build component of the final

design was determined by measurement of the CATIA V.5 models in the as drawn

configuration i.e. as geometric blanks sized to the generalised IML and then

subtracting the delta percentage difference between the blank weight and the as drawn

weight of a nominally sized structure as shown in figures E-15 and 16. The densities

for the structural materials used were taken from references 33 which were listed in

the reference section 6 on page 207 of this thesis.

The densities applied were as follows:-

2024-T351 Aluminium plate = 2768 kg/m3 (0.1 lb/in

3)

2104-T6 Aluminium forgings = 2768 kg/m3 (0.1 lb/in

3)

BMI / graphite = 1522.4 kg/m3 (0.055 lb/in

3)

6Al-4V Titanium plate = 4428.8 kg/m3 (0.16 lb/in

3)

6Al-6V Titanium forgings = 4539.52 kg/m3 (0.614 lb/in

3)

Al alloy 2024-T351 plate frame :-

8mm flanges: 4mm web: 4mm

stiffener thicknesses.

Ti alloy 6Al-6V-2Sn forging frame :-

8mm flanges: 3mm web: 4mm


BMI carbon RTM frame :- 7mm

flanges: 4mm web: 25mm top hat


Weight comparison.

As drawn Block Frame with weight

measured in Al: Ti: and BMI carbon.

As drawn Design intent Flyaway Part

Al Block frame = 183.3kg Pocketed frame = 20.6kg Actual weight % = 11

Ti Block frame = 300.5kg Pocketed frame = 27.4kg Actual weight % = 9.1

BMI Block frame = 100.8kg BMI RTM frame = 19.4kg Actual weight% = 19.2

Figure E-15:- Weight comparison between the representative „As drawn frames‟

and „Notional thickness detailed design intent frames‟ to determine the

percentage of the as drawn frame weight an actual „Flyaway frame‟ would equal.



246

Al alloy 2024-T351 plate Keel :- 4mm

flange and 4mm web thicknessBMI Carbon RTM Keel :- 4mm

flange and 4mm web thickness

Ti alloy 6Al-4V plate keel :- 4mm

flange and 4mm web thickness

As drawn Block Frame with weight

measured in Al: Ti: and BMI carbon.

Weight comparison.

As drawn Design intent Flyaway Part

Al Block keel = 369.7kg Pocketed keel = 46.2kg Actual weight % = 12.5

Ti Block keel = 591.4kg Pocketed keel = 73.9kg Actual weight % = 12.5

BMI Block keel = 203.3kg BMI RTM keel = 25.4kg Actual weight% = 12.5

Figure E-16:- Weight comparison between the representative „As drawn keels‟

and „Notional thickness detailed design intent keels‟ to determine the percentage

of the as drawn keel weight an actual „Flyaway keel‟ would equal.

Table E.1:- Ti Wing & Empennage structural weight measurements.

Wing component Weight in kg

Port Wing boundary structure 281.12

Stbd Wing boundary structure 281.12

Port Wing core structure 305.28

Stbd Wing core structure 305.28

Port leading edge ribs 10.93

Stbd leading edge ribs 10.93

Total wing weight 1,194.64

Empennage components Weight in kg

Port / Stbd boundary structure 91.08

Port / Stbd core structure 69.02

Total empennage weight 160.10



247

Table E.2:- Forward fuselage structural weight measurements.

Component Frame Station in m*

FB-24 Weight in kg A-24 Weight in kg

Chevron Frame 11.15 11.15

Frame 2.51 13.37 13.37

Frame 2.92 8.87 8.87

Frame 3.35 7.85 7.85

Frame 3.71 10.48 10.48

Frame 4.09 17.30 17.30

Frame 4.46 20.67 20.67

Frame 4.88 17.42 32.26

Frame 5.35 25.52 44.00

Frame 5.70 44.08 59.58

Frame 6.12 50.36 62.95

Frame 6.43 64.60 64.60

Keels (Port / Stbd) 41.80 41.80

Chine longerons (Port / Stbd)

4.14 4.14

Outboard Longerons (Port / Stbd)

1.78 1.78

Canopy landings (Port / Stbd)

2.36 2.36

Hatch landings (Port / Stbd)

1.16 N/A

Frame 6.43 stiffeners 3.32 N/A

Hatch 92.77

Total weight 439.00 403.16

* SEE FIGURE E-2 FOR FRAME STATION LOCATIONS.



248

Table E.3:- Centre and Aft fuselage structural weight measurements.

Centre Fuselage Component*

Weight in kg

Aft Fuselage Component*

Weight in kg

Frame 7.12 106.45 Frame 14.39 79.65

Frame 8.38 76.98 Frame 14.92 74.90

Frame 9.52 80.62 Frame 15.41 67.16

Frame 10.62 79.42 Frame 15.89 58.08

Frame 11.60 89.87 Frame 16.37 46.36

Frame 12.58 88.29 Frame 17.08 (x2) 11.88

Frame 13.15 87.85 Frame 17.94 (x2) 5.48

Frame 13.97 113.42 Kicked Keel (Port /

Stbd) 43.20

Lower Keels (Port / Stbd)

98.18 Shoulder longerons

(Port / Stbd) 20.94

Shoulder longerons (Port / Stbd)

21.74 Lower Longerons

(Port / Stbd) 38.76

Mid aft Keels (Port / Stbd)

65.14 Chine longerons

(Port / Stbd) 11.04

Total weight 907.96 Keel beam

(Port / Stbd) 80.8

I/B Fairing chine (Port / Stbd)

3.96

Boom closure (Port / Stbd)

0.74

Total weight 636.01

*COMMON STRUCTURAL COMPONENTS FRAME STATIONS IN

METRES SEE FIGURES E-6 AND E-7 FOR FRAME STATION LOCATIONS.

The weights measured for the aircraft skins were taken directly from solid models

created by the author (as were all models contained herein). The wing skin lay up was

a simplified proportional representation without skin ramps and consisted of stepped

wing skins from 20mm at the root to 4mm at the tip as shown in figure E-17. The

fuselage skins were constant 5.08mm OML off-set solids without pad-up areas, as

shown in figure E-18, and the same method was used to generate the ruddervator

skins. The weights of these skins are given in table E-4 below.

The total measured weights of the major structural components are compared with the

estimated component weights in lbs and kgs in table E-5 below, feeding into the

definitive weight statement in table E-6 below.



249

Figure E-17:- Weight of wing skins to a notional balanced ply lay-up

configuration.

Figure E-18:- Weight of fuselage skins to a notional balanced ply lay-up

configuration.



250

Table E.4:- Major component skin weight measurements.

Skin component Weight in kg

Port Wing skin 944.00

Stbd Wing skin 944.00

Forward fuselage skin 185.12

Centre fuselage skin 305.52

Aft fuselage skin 123.41

Port ruddervator skin 72.50

Stbd ruddervator skin 72.50

Total component weight Weight in kg

Total Wing skin weight 1,888.00

Total Fuselage skin weight 614.05

Total Ruddervator skin weight 145.00

Total skin weight 2,647.05

Table E.5:- Major component weight measurements compared with

AeroDYNAMICTM

estimated values.

Major airframe component

Measured weight in kg / lb

Estimated weight in kg / lb

Wings 3,082.64 / 6,796.06 1,232.40 / 2,716.90

FB-24 Fuselage 2,597.02 / 5,725.45 2,837.60 / 6,255.70

A-24 Fuselage 2,561.18 / 5,646.44 2,837.60 / 6,255.70

Verticals 497.96 / 1,097.8 323.2 / 712

Undercarriage Main 583.64 / 1,286.70 590.40 / 1,301.70

Undercarriage Nose 243.40 / 536.60 In Main U/C value

Engine weight 3,402.00 / 7,500.00 3,402.00 / 7,500.00

GTOW FB-24 19,989.24 / 44,068.73 22,327.4 / 49,223.5

GTOW A-24 21,771.86 / 47,998.73 22,327.4 / 49,223.5

Landing weight with 10% fuel FB-24

10,406.66 / 22,942.76 11,208.00 / 24,709.50

Landing weight with 10% fuel A-24

10,484.83 / 23,115.09 11,208.00 / 24,709.50



251

The measured structural weights to notional sizings are higher for the wings and

ruddervators, than the AeroDYNAMIC values and yet significantly lower for the

fuselage, this was attributed to the generalised thick skins used for the wings and

reduction in internal structure in the fuselage. This leads to the definitive weight

statement below.

Table E.6:- Definitive as designed weight measurements statement.

Major aircraft component Measured weight in kg / lb

Wings 3,082.64 / 6,796.06

FB-24 Fuselage 2,597.02 / 5,725.45

A-24 Fuselage 2,561.18 / 5,646.44

Verticals 497.96 / 1,097.8

Undercarriage Main 583.64 / 1,286.70

Undercarriage Nose 243.40 / 536.60

Propulsion weight 3,402.00 / 7,500.00

Equipment weight* 2,233.81 / 4,924.70

Design empty weight FB-24 12,640.47 / 27,867.46

Design empty weight A-24 12,604.63 / 27,788.45

Useful load weight* F-24 11,484.96 / 25,320.00

Useful load weight* A-24 13,203.62 / 29,109.00

GTOW FB-24 24,125.43 / 53,187.46

GTOW A-24 25,808.25 / 56,897.45

Landing weight with 10% fuel FB-24

13,598.73 / 29,980.06

Landing weight with 10% fuel A-24

13,744.74 / 30,301.95

*Employs releasable figures from F-35 Joint Project Office (no individual

weights to be identified ITAR restrictions).

N.B.:- Fuel capacity of the F-24 = 9,582.59kg = 21,126.00lb

N.B.:- Fuel capacity of the A-24 = 11,401.04kg = 25,135lb

Equipment: - Flight controls: Instruments: Hydraulics: Electrical: Avionics: GFE

(ejection seats): Air conditioning: Handling gear: Engine heat shield: etc.

Useful load: - Crew: Fuel: Oil: Stores (weapons and defensive aids Chaff / flares):

mission critical devices NVG‟s: etc.



252

This concluded the structure and weight study for the F-24 and A-24 PWSC

designs.

E-2 The Subsonic and Supersonic range and endurance values for the FB-24 and

A-24 aircraft NB-2 configurations.

E-2.1:- The maximum endurance and range for A-24 at BCM at 34,400ft.

For the A-24 configuration the maximum endurance is achieved at the speed for

(L/D)max which can be determined by first calculating the required value of CL, then

solving for the speed required to achieve L=W at that CL therefore for A-24:-

CL = CDo / k = 0.014361 / 0.167051 =0.293


2



V = 2q / = 2 x (188.131lb/ft2) / 0.000767 slug/ft

3 = 700.4ft/s

for maximum endurance. Note that this is only the initial velocity for maximum









that for W2

W2 = W1 – Wf = 56,897.45lb – 25,135lb = 31,762.45lb

E = 1/ ct CL / CD ln (W1 / W2) = 1 / 0.78 (10.2) ln (56,897.45 / 31,762.45) = 7.62h

Therefore for A-24 the maximum endurance at 34,400ft at Best Cruise Mach

BCM is 7.62hours.

Similarly the value for maximum range is obtained by solving equation 3.36 for CL

and equation 3.37 for q:-



253

CDo = 3k CL2, CL = CDo / 3k = 0.014361 / 3(0.167051) = 0.239

q = W / CLS = 56,897.45 / 0.239(893ft2) = 217.45lb/ft

2

V = 2q / = 2 x (217.45lb/ft2) / 0.000767 slug/ft

3 = 753.00ft/s




CD = 0.014361 + 0.167051 CL2 = 0.014361 + 0.167051(0.239)

2 = 0.0239

R = 2/S / ct CL2 / CD (W1

1/2 – W2

1/2)

R= 2 / (0.000767 slug/ft3) (893ft

2) 2 / 0.78 (0.239)

1/2 / 0.0239

x ((56,897.45lb)1/2

-(31,762.45lb)1/2

)

R = 15.315 (ft2 / lbs

2) (2.564h) (20.45) (12.568lb

1/2) = 2,263.97miles


Therefore the A-24 would have a range value of R = 1,457.97nmiles.

E-2.2:- Supersonic wave drag and critical Mach number analysis.

This performance appeared to be suspiciously low and can be attributed to the high

drag values based on the crude modelling in AeroDYNAMIC V2.08. For supersonic

range and performance values were based on the CATIA V5 model geometry to

predict MCrit C D wave and k1 using the methods in reference 22 as follows:-

MCrit (unswept) = 1.0 – 0.065 [100 (tmax / c)] 0.6

= 1.0 – 0.065 (5)0.6

= 0.82

MCrit = 1.0 – cos 0.6

0.25 (1.0 - MCrit (unswept))

= 1.0 – cos 0.6

43.90 (1 – 0.82) = 0.87

MCDo max = 1 / cos 0.2

LE = 1 / cos 0.2

520 = 1.0

C D wave = 4.5 / S (Amax / l) 2 EWD (0.74 + 0.37 cos LE) [1 - 0.3 M - MCDo max]

C D wave = 0.0158 x (24.46 / 60) 2 x 1.355 x 0.379 = 0.0134 at Mach 1.6

k1 = [AR (M2 – 1) / (4AR M

2 – 1) – 2] cos LE



254

k1 = [2.27 (1.62 – 1) / (4 x 2.27 1.6 – 1) – 2] cos 52

0 = 0.208

The parasitic drag was estimated using the equivalent skin friction coefficient of 0

0025 (representative of a smooth fast jet transport aircraft) from reference 1 and the

CATIA V5 surface wetted area of 161m2 = 1,732ft

2 therefore the parasitic drag value

was:-

C D min = 0.0025 (1,732 / 893) = 0.00485

This gave a total clean aircraft zero lift coefficient of 0.01825 in supersonic cruise

condition

E-2.3:- The maximum endurance and range for A-24 at Mach 1.6 and 45000ft.

Maximum supersonic endurance at Mach 1.6 (L/D)max was determined by first

calculating the value of CL, then solving for the speed required to achieve L=W at that

CL therefore for A-24:-

CL = CDo / k = 0.01825 / 0.208 = 0.6495

L = W = CLqS, q = W / CLS = 56,897.45 / 0.6495 (893ft2) = 98.10b/ft

2



V = 2q / = 2 x (98.10lb/ft2) / 0.000462 slug/ft

3 = 651.67ft/s

For maximum endurance: Note that this is only the initial velocity for maximum








Then the endurance can be calculated using equation 3.36 with W1 = 48,200 and that

for W2

W2 = W1 – Wf = 56,897.45lb – 25,135lb = 31,762.45lb

E = 1/ ct CL / CD ln (W1 / W2) = 1 / 1.45 (8.3) ln (56,897.45 / 31,762.45) = 3.34h



255

Therefore for A-24 the maximum endurance at 45,000ft at Mach 1.6 is

3.34hours.

Similarly the maximum range value for the FB-24 is obtained by solving equation

3.36 for CL and equation 3.37 for q:-

CDo = 3k CL2, CL = CDo / 3k = 0.01825 / 3(0.208) = 0.1710

q = W / CLS = 56,897.45 / 0.1710(893ft2) = 372.601b/ft

2

V = 2q / = 2 x (315.645lb/ft2) / 0.000462 slug/ft

3 = 1,168.94ft/s




CD = 0.01825 + 0.208 CL2 = 0.01825 + 0.208(0.1710)

2 = 0.0243

R = 2/S / ct CL2 / CD (W1

1/2 – W2

1/2)

R= 2 / (0.000462 slug/ft3) (893ft

2) 2 / 1.45 (0.1710)

1/2 / 0.0243

x ((56,897.45lb)1/2

-(31,762.45lb)1/2

)

R = 4.84 (ft2 / lbs

2) (1.378h) (17.02) (12.568lb

1/2) = 648.48miles


Therefore the value for R = 383.72nmiles.

E-2.4:- The maximum endurance and range for FB-24 at Mach 1.6 and 45000ft.

Maximum supersonic endurance at Mach 1.6 (L/D)max was determined by first

calculating the value of CL, then solving for the speed required to achieve L=W at that

CL therefore for FB-24:-

CL = CDo / k = 0.046832 / 0.349518 = 0.6495


2



V = 2q / = 2 x (91.70lb/ft2) / 0.000462 slug/ft

3 = 630.05ft/s



256

For maximum endurance: Note that this is only the initial velocity for maximum









that for W2

W2 = W1 – Wf = 53,187.46lb – 21,126lb = 32,061.46lb

E = 1/ ct CL / CD ln (W1 / W2) = 1 / 1.45 (8.3) ln (53,187.46 / 32,061.46) = 2.90h

Therefore for FB-24 the maximum endurance at 45,000ft at Mach 1.6 is

2.90hours.

Similarly the maximum range value for the FB-24 is obtained by solving equation

3.36 for CL and equation 3.37 for q:-

CDo = 3k CL2, CL = CDo / 3k = 0.01825 / 3(0.208) = 0.1710

q = W / CLS = 53,187.46 / 0.1710(893ft2) = 348.31lb/ft

2

V = 2q / = 2 x (348.31lb/ft2) / 0.000462 slug/ft

3 = 1,227.93ft/s




CD = 0.01825 + 0.208 CL2 = 0.01825 + 0.208(0.1710)

2 = 0.0243

R = 2/S / ct CL2 / CD (W1

1/2 – W2

1/2)

R= 2 / (0.000462 slug/ft3) (893ft

2) 2 / 1.45 (0.1710)

1/2 / 0.0243

x ((53,187.46lb)1/2

-(32,061.46lb)1/2

)

R = 4.84 (ft2 / lbs

2) (1.378h) (17.02) (10.563lb

1/2) = 541.86miles



257


Therefore the value for R = 320.63nmiles.

*Therefore the maximum range and endurance of the FB-24 two place manned

aircraft was:-

At 34,400ft and Mach 0.85: -

Endurance = 7.03 hours

Range = 1,282.54nmiles

At 45,000ft and Mach 1.6: -

With A/B Endurance = 2.90hours and without A/B Endurance = 5.80hours.

With A/B Range = 320.63nmiles and without A/B Endurance = 640nmiles.

* Therefore the maximum range and endurance of the A-24 UCAV unmanned

aircraft was:-

At 34,400ft and Mach 0.85: -

Endurance = 7.62hours

Range = 1,457.97nmiles

At 45,000ft and Mach 1.6: -

With A/B Endurance = 3.34hours and without A/B Endurance = 6.68hours.

With A/B Range = 383.72nmiles and without A/B Endurance = 767.44nmiles.

N.B. Mach 1.6 values for both aircraft are given with and without afterburner

engaged the endurance and range of the YF-120 in dry supersonic flight

condition is twice that with afterburner with 80% of the thrust based on USAF

released data reference 26.

E-3:- A description of the AeroDYNAMICTM

V2.08 toolset analysis codes.

As explained at length above the Jet306 tool could not be used for this conceptual

design study due to numerous run – time errors, and a replacement was not available

in the required time frame therefore the older less sophisticated AeroDYNAMICTM

V2.08 tool was used and it is this which is described below.

A complete aircraft will frequently generate significantly more lift than its wing

alone. AeroDYNAMICTM

estimated the whole aircraft‟s lift by summing the lift

contributions of its various components. The methodology employed and described

below was a simple means of making an initial estimate of an aircraft‟s aerodynamic

capabilities and was suitable for use in the early conceptual phases of design, and was

deemed by the author as applicable to this design proposal submission.



258

E-3.1:- Lift analysis.

For most aircraft including the FB-24 / A-24 the majority of the lift is generated by

the wing with the pitch and trim surfaces contributing additional lift.

AeroDYNAMICTM

used the non-elliptical finite wing expression to determine the lift

curve slope per degree of angle of attack CL and CLt for the wing and the pitch and

trim surfaces as shown as equations E.1 and E.3 below. The Span efficiency factor

was determined for the wing and the pitch and trim surfaces using empirical

expression for e shown in equations E.2 and E.4 below.

Wing

0.052019 per degree

where:

0.659126

Pitch Trim Surface

0.052019 per degree

where:

0.659126

Cc

c

e AR

L

l

l

157 3.

eAR AR t

2

2 4 12 2( tan )max

t

l

l

tL

ARe

c

cC

3.571

eAR AR t

2

2 4 12 2( tan )max

Equation block 3.1(a): - AeroDYNAMICTM

Aerodynamic lift analysis codes.

Where:-

AR = aspect ratio of the wing or pitch and trim surface:

t max = the sweep angle of the line connecting the point of maximum thickness

on each airfoil section of each wing or pitch and trim surface:

cl = lift coefficients at the angle of attack

5.73 = section lift – curve slope, per radian.

The data generated from this analysis was used by AeroDYNAMICTM

in the

generation of the following charts: - 1: 2(a) and,2(b): 3: 4: 8: 9(a) and (b): 10: and 11

shown in pages 136 to 142 inclusive.

(E.1)

(E.2)

(E.3)

(E.4)



259

An aircraft fuselage is usually long and slender and therefore dose not produce much

lift. In the region of the horizontal lifting surfaces, however, the lift being generated

by those surfaces carries over the fuselage. In order to model this effect the wing is

treated as if it extends all the way through the fuselage without any changes in airfoil,

sweep, or taper. In fact, the fuselage shape is significantly different from the wings

airfoil shape and may be less effective at generating lift. However, because the

fuselage lifting area is generally larger than the portion of the wing in the fuselage, the

two effects may be treated as cancelling each other out, at least for early conceptual

design. For a design with strakes or LEXs, the effect is modelled in

AeroDYNAMICTM

as shown in equations E.5 and E.6 below (equation block 3.1(b),

based on wind tunnel testing, where S strake. includes only the exposed area of the

strake, and not any portion inside the fuselage. Because an angle of attack = 150 is

usually the maximum useable equations E.5 and E.6 are adequate for the useable

range (reference 22).

Equation block 3.1(b): - AeroDYNAMIC

TM Aerodynamic lift analysis codes

continued.

To determine the horizontal tail or canard surfaces contribution to the whole aircraft

lift curve slope it is first necessary to determine the rate at which downwash (or

upwash, as appropriate) changes with changing aircraft angle of attack.

AeroDYNAMICTM

estimated the rate of change of downwash with angle of attack

using the empirical curve fit given above as equation E.7 in equation block 3.1(b),

(based on wind – tunnel testing rather than theory). Where cavg is the mean geometric

chord of the wing: lh is the distance from the quarter-chord point of the average chord

of the main wing to the quarter – average – chord point on the horizontal surface, as

shown below in figure E-19: Zh is the vertical distance of the horizontal surface above

the plane of the main wing, as shown in figure E-19: and CL is in degrees.

Whole Aircraft

0.065639 +

where:

0.055487

and:

0.457872

C CL L ( ) ( )whole aircraft with strake CLt

1

S

S

t

C CS S

SL L

strake

( ) ( )with strake without strake

21 10 3

71

0 725

o C

AR

c

l

z

b

L avg

h

h .

zh

lh

.25 croot

.25 croot

lc

.25 croot

(E.5)

(E.6)

(E.7)



260

Once the value of / has been predicted, the horizontal surfaces contribution to

the aircrafts CL can be approximated as: -

CL (due to horizontal tail) = CLt (1 - / St / S E.8

were the subscript t denotes parameters for the horizontal tail.

The common CL (due to horizontal tail) values vary from almost zero to 35% or

more of CL.

Figure E-19:- Aircraft geometry for downwash predictions.

For horizontal surfaces ahead of the wing, also known as canards AeroDYNAMICTM

another empirical equation based on wind – tunnel experimentation for predicting the

rate of change of upwash with angle of attack, for wings with a quarter chord sweep

angle 0.25 < 350 expression used was: -

u / = (0.3AR0.3

– 0.33) (lc / c)-(1.04+6AR^-1.7)

E.9

Where u was the upwash angle and lc was the distance from the wings quarter –

chord to the canards quarter – chord as shown in figure E-19 above.

Once u / had been estimated, the canards contribution to the aircrafts CL was

approximated as:-

CL (due to canard) = CLc (1 - u / Sc / S E.10

Where the subscript c identifies quantities related to the canard. Contributions of

canards to the total aircraft lift curve slope are typically larger than those for the

horizontal tails. This is partly due to the canard being in an upwash field rather than

the downwash field surrounding most horizontal tails. Once the contributions of

canards or horizontal tails have been estimated, the whole aircraft lift curve slope is

generated from:-

CL (whole aircraft) = CL (wing + body + strakes) + CL (due to horizontals) +

CL (due to canard) E.11

zh

lh

.25 croot

.25 croot

lc

.25 croot

0.25croot

0.25croot

0.25croot



261

E-3.2:- Drag and Critical Mach number analysis.

The drag for the complete aircraft is identified as either parasite drag or drag due to

lift. Parasite drag is all of the drag on the aircraft when it is not generating lift. This

includes both skin – friction and pressure drag, as well as several additional types of

zero-lift drag that are associated with the complete aircraft configuration. The drag

due to lift includes all types of drag that depend on the amount of lift that the aircraft

is producing which includes induced drag due to downwash, the pressure drag, which

increases with lift due to forward movement of the separation point, induced and

pressure drag from canards and horizontal tails, and additional drag, e.g. vortex drag

due to leading-edge vortices on strakes and highly swept wings. All of the above

types of drag can be approximated by the simple expression for the drag coefficient

given below as equation E.5 which was used by AeroDYNAMICTM

for Drag Polar

determination, shown in figures 1 and 8 on pages 136 and 140 respectively.

Drag Polar:

Mach CDmin CDo k1 k2

0.1 0.01792 0.01800 0.1168 -0.0061

0.872737 0.01792 0.01800 0.1168 -0.0061

1.054749 0.05253 0.05261 0.2419 -0.0046

1.6 0.04486 0.04494 0.3045 0.0000

2 0.04244 0.04251 0.3670 0.0000

0.872737 M crit = 1.0 - 0.065 cos0.6

LE

add B-1 data to WIM-13

C C k C k CD Do

L L 1

2

2

100

0 6t

c

max

.

Equation block 3.2(a): - AeroDYNAMIC

TM drag and critical Mach number

analysis codes.

Equation E.12 where:-

k1 = 1 / ( e0 AR) (subsonic): - e0 = Oswald‟s efficiency factor: AR = Aspect

Ratio = b2 / S were b

2 is the wing span squared and S is the wing area:

k1 = AR (M2 – 1) / (4ARM

2 – 1) – 2 (supersonic): - AR = Aspect Ratio: M

= Mach number:

CDo = the parasite drag coefficient:

k2 = - 2 k1 CL min D (negative value subsonic and zero value supersonic):

CL min D = the CL value for which CD is a minimum:

k1CL2 = induced drag.

(E.12)

(E.13)

0.6



262

The value of k2 is chosen to allow modelling of wings with airfoils that generate

minimum drag at some nonzero value of lift.

Parasite drag: - The first step in determining Parasite drag in AeroDYNAMICTM

was

to determine the wetted area by approximate the aircraft as a set of simple geometric

shapes, as described in section 3.3.2 page 134. The equations for the surface areas of

these simple shapes were well known, and by taking care to avoid counting areas

where two shapes touch, it was relatively easy to determine Swet the shapes available

to define the fuselage were: - cylinders: half - cylinders: circular: elliptical:

rectangular: cone: and half - cone. The surface areas of these shapes did not include

ends which butted up against another shape because such ends would not be wetted.

When the longitudinal flat face of a half cylinder or half - cone touches another body,

twice the surface area of that face was subtracted, because it and an equal area of the

other body were in contact with each other, and were therefore not wetted. The

margin of error within AeroDYNAMICTM

for relatively simple geometry aircraft e.g.

F-15, F-16 was 5% depending on the number of data points and hence the fidelity of

the model.

The skin - friction drag on a complete aircraft configuration is generally much greater

than on the wing alone because the wetted area is much greater. To make the initial

estimate of subsonic parasite drag AeroDYNAMICTM

used the concept of equivalent

skin – friction drag coefficient, Cfe employing data values from large numbers of

similar aircraft types (see table 4.1 in equation block 3.2(b) bellow), which is as

follows: -

Cfe = CDo S / Swet E.14

These values are based on historical data and are a function of such diverse factors as:

- aircraft skin materials, shape, paint, typical flight Reynolds number, number of

additional air scoops for ventilation, type size, number, and location of engine intakes,

and attention to detail in sealing doors, control surface gaps, etc. Using Cfe to predict

CDo for an aircraft that generates minimum drag when it is generating zero lift only

requires selecting a Cfe for the appropriate category of aircraft and estimating the total

wetted area of the aircraft concept as described above. The value of CDo could then be

obtained by solving

CDo = Cfe S / Swet E.15

As shown in equation block 3.2(b).

Drag due to lift:- AeroDYNAMICTM

predicted Oswald‟s efficiency factor with a

curve fit of wind tunnel data for a variety of wing and wing – body combinations and

the equation for this curve is shown as equation E.16 in equation block 3.2(b). Note

that increasing wing sweep decreases the value of eo. This is because for high aspect

ratio wings, that part of the airfoil profile drag that varies with lift is a larger part of

the total drag due to lift that eo must model.



263

subsonic, supersonic

0.017921 0.0037 from Table 4.1

1.054749

0.116774 k 1 = 1/( e o AR) subsonic

where:

0.908619

supersonic

Table 4.1 Common C fe Values

-0.006074 Type C fe

subsonic Jet Bomber and Civil Transport 0.003

Military Jet Transport 0.0035

0.026009 Air Force Jet Fighter 0.0035

Carrier-Based Navy Jet Fighter 0.004

supersonic Supersonic Cruise Aircraft 0.0025

Light Single Propeller Aircraft 0.0055

Light Twin Propeller Aircraft 0.0045

clean 1.0502 Assume max = 15 degrees Propeller Seaplane 0.0065

Jet Seaplane 0.004

takeoff 1.2538 DCLmax =2/3 of DCLmax for Landing, CLmax = CLmax clean + DCLmax

landing 1.3587 DCLmax = 0.5 * S h /S * l h /(0.5*c root ) = max trimmable DCL

e ARo LE 4 61 1 0 045 310 68 0 15. ( . )(cos ) .. .

k k CLminD2 12

CS

A

lE M MD

maxWD LE C maxwave Do

4 50 74 0 37 1 3

2.

. . cos .

MC

LEDo max .cos

1

0 2

k

AR M

AR MLE1

2

2

1

4 1 2

cos

2

0L

LminD

L CC

S

SCC wet

fD emin

C C k CD D Lo min minD 1

2

efC

100

0 6t

c

max

.

wavesubsonico DDD CCC 0

02 k

C C CLmax L amax L max L ( )0

Equation block 3.2(b): - AeroDYNAMIC

TM drag and critical Mach number

analysis codes.

Effects of Camber:- AeroDYNAMICTM

determined profile drag resulting from:-

cambered airfoils: fuselage shape and orientation, which can result in minimum drag

being generated at a positive (non zero) value of lift coefficient, by using the

additional k2 CL term in equation E.12 in equation block 3.2(a) above. For example,

the minimum drag coefficient for an aircraft occurs at a lift coefficient signified by the

symbol CL min D then the necessary value of k2 is given by: -

k2 = - 2 k1 CL min D E.18

The value of CL min D is determined from an approximation which assumes that the

airfoil generates minimum drag when it is at zero angle of attack and that the effect of

induced drag is to move CL min D to a value half way between zero and the value of

CL when = 0. The value of CL when = 0 being given by: -

CL min D = CLa () = CLa (- L = 0) E.19

Because = - L = 0 and = 0 equation E.19 becomes

(E.16)

(E.17)



264

CL min D = CLa (- L = 0 / 2) E.20

The equation is shown above in AeroDYNAMICTM

equation block 3.2(b). This value

of CL min D was used for the entire aircraft. This is done because the assumption was

made that aircraft fuselage and strakes etc were designed so that they also had their

minimum drag at the angle of attack that places the wing at its CL min D. When this is

done, the minimum value of C D which is given the symbol C D min must not be any

lower than C Do predicted by equation E.15 above. Recalling that C Do is the aircrafts

zero – lift drag coefficient. For aircraft with minimum drag at non zero lift, this

resulted in the following revised predictions: -

CD min = Cfe S / Swet E.21

C Do = CD min + k1 C 2L min D E.22

Critical Mach number:- AeroDYNAMICTM

uses equation E.13 in equation block

3.2(a) to determine the critical mach number of the configuration under analysis, this

substitutes swept wing chord equation (E.23) into the curve fit of Mcrit data for the

NACA 64-seies airfoils (used when actual airfoil data is not available in conceptual

design analysis).

Therefore: -

( t max / c ) (swept wing) = ( cos LE ) ( t max / c ) (unswept wing) E.23

Becomes: -

Mcrit = 1.0 – 0.065 cos0.6

LE [100 ( t max / c )]0.6

E.24

There by producing an expression the critical Mach number for 3-D swept wings, for

the unswept wings AeroDYNAMICTM

uses:-

Mcrit = 1.0 – cos0.6

LE (1.0 – Mcrit (unswept)) E.25

For tapered wings as in the planforms types evaluated in this conceptual design study,

AeroDYNAMICTM

uses 0.25c, the sweep angle of the line connecting the quarter-

chord points of the wings airfoils and using the maximum value of (t max /c) on the

wing i.e.:-

Mcrit = 1.0 – cos0.6

0.25c (1.0 – Mcrit (unswept)) E.26

*See also Reference 22: -page 123: Introduction to Aeronautics A Design

Perspective: by Brandt S. A: Stiles R. J: Bertin J. J: Whitford R.: Published by AIAA:

1997



265

Supersonic Drag Due to Lift:- AeroDYNAMICTM

uses equation E.17 in equation

block 3.2(b) to calculate the supersonic value of k1. At supersonic speeds all airfoils,

regardless of shape generate zero lift at zero angle of attack. Practical supersonic

airfoil shapes also generate minimum drag at zero angle of attack, and so in the

supersonic regime, k2 = 0. The supersonic value of k1.is given by:-

k1.= [AR (M2 – 1) / (4AR M

2 – 1) – 2] cos LE E.27

Where:-

AR = Aspect Ratio:

M2 = Mach number:

LE = Sweep angle of the leading edge.

For a well designed supersonic aircraft, the transition from subsonic to supersonic

values of k1 and k2 is gradual, so that the variation of these parameters through the

transonic regime can be approximated with a smooth curve, as shown in charts 1 and

8 in the main body of this report on pages 136 and 140 respectively for the NB1 and

NB2 configurations of the F-24 / A-24 configurations studied in this report.

Total Drag:- In summary, the total drag on an aircraft is the sum of the profile drag

(the subsonic drag not due to lift), wave drag, and the drag due to lift or induced

drag:-

CDo = CDp + CD wave E.28

And:-

CD = CDo + k1 CL2 + k2 CL E.29

And AeroDYNAMICTM

uses these equations as shown in equation blocks 3.2(a) and

(b) above to determine total drag for the configuration under study, for the NB1 and

NB2 configurations investigated in this report the results are shown graphically as

charts 1(a) / (b) and 8(a) / (b) respectively on pages 136 and 140 of this report.

This concludes the overview of the aerodynamic codes and predictions used by

AeroDYNAMICTM

analysis tool, as can be seen this is a basic conceptual design tool

employing Microsoft Excel spreadsheets to process classical aerodynamic analysis

equations and although not very sophisticated it is the equal of a similar older MS.Dos

tool developed by Daniel Raymer and is deemed adequate for use by the United States

Airforce Academy as an academic design analysis tool. Unlike some conceptual

design approaches the user dose have to generate a concept rather than pages of

algebra, which is one reason the design approach taken in this report may appear odd

to some academics.



266

E-4:- The application of AeroDYNAMICTM

V2.08 toolset to the NB2 / F-24 and

A-24 configuration.

The analysis of the F-24 / A-24 final airframe configuration conducted using

AeroDYNAMICTM

V 2.08 resulting in the charts 8 through 14 pages 140 – 143 used

applied the aerodynamic analysis methods and codes described above to the aircraft

configuration data inputs as described in the main report and below to generate the

following intermediate analysis stages.

Figure E-20:-Lifing surface analysis data for the common F-24 / A-24

configuration.

This data was used to calculate the aspect ratio AR of the wing and the ruddervators

and from this the value of e was determined for both surfaces and the two

dimensional lift curve slope for the NACA 64-0006 was approximated allowing CL

to be determined from equation E.1 for each surface as presented in the previous

section above. This configuration had no strakes so the CL(with strakes) term made no

contribution to the whole aircraft lift coefficient. The distance from the quarter chord

of the main wings mean chord to the same point on the aircrafts ruddervator l h the

wings taper ratio and the distance of the ruddervators above the wing Zh was used

to determine from equation E.7 described in the previous section above.

Finally using the wing and ruddervator areas S and S r the whole aircraft coefficient

CL(whole aircraft) was determined from equation E.5 as described in the previous

section above.



267

Form these calculations the lift curve and aerodynamic analysis data shown below in

tables E-7 and E-8 were produced.

Table E-7:- Lift curve analysis data for the F-24 / A-24.

alpha CL CLalpha CLmax CLmaxTO CLmaxL

1 -2 -0.090175 0.0451 0.68 1.04 1.23

2 -0 0 [alpha at L = 0] Sref Swet Amax Length

3 3.5 0.157806 906.98 2835.41 72.56 60.00

4 7 0.315611

5 10.5 0.473417

6 14 0.631223

7 16 0.6763 [alpha at Clmax]

8 18 0.631223

Table E-8:- Aerodynamic analysis data for the F-24 / A-24.

Horiz Surf AR Taper Sweep S Section clAlpha e k

1 2.2327 0.07094 52 906.975 NACA0006 2.535512 0.853437 0.167051 L= 60

2 2.5002 0.23283 52 332.20814 NACA0006 2.567374 0.826795 0.153983 sWet= 2835.41321

3 f= 11.3416528

4 cfe= 0.004

5 cdo= 0.01250492

6 clat= 2.5833153

7 clMax= 0.67631036

8 clMaxTO= 1.04220329

9 clMaxL= 1.22514975

10 sRef= 906.975

11 clopt= 0

12 amax= 75.9676318

13 amax - A0= 72.557718

14 vol= 2653.8041

15 clAlpha= 2.583315

16 fuelVolume = 0

17 systemVolume = 0

18 expPayloadVolume = 0

19 seatVolume = 9

20 gearVolume = 86.5956engineVolume = 200.4529

The drag of the final F-24 / A-24 common configuration was calculated as follows

from the input data starting with the parasite drag. The first step in determining

Parasite drag in AeroDYNAMICTM

was to determine the wetted area by approximate

the aircraft as a set of simple geometric shapes, as described in section 3.3.2 page 134,

and shown below in figure E-21. The equations for the surface areas of these simple

shapes were well known, and by taking care to avoid counting areas where two shapes

touch, it was relatively easy to determine Swet the shapes available to define the

fuselage were: - cylinders: half - cylinders: circular: elliptical: rectangular: cone: and

half - cone. The surface areas of these shapes did not include ends which butted up

against another shape because such ends would not be wetted. When the longitudinal

flat face of a half cylinder or half - cone touches another body, twice the surface area

of that face was subtracted, because it and an equal area of the other body were in

contact with each other, and were therefore not wetted.



268

The margin of error within AeroDYNAMICTM

for relatively simple geometry aircraft

e.g. F-15, F-16 was 5% depending on the number of data points and hence the fidelity

of the model.

CONE

HALF CONE

HALF CYLINDER

HALF CONE

SURFACE #1

SURFACE #2

SURFACE #4

SURFACE #3

SURFACE #6

SURFACE #5

SURFACE #8

SURFACE #9

OVAL CYLINDER #1

OVAL CYLINDER #2

OVAL CYLINDER #3

OVAL CYLINDER #4

CYLINDER #1

Figure E-21:- F-24 / A-24 configuration geometry approximation by simple

shapes.

The skin - friction drag on a complete aircraft configuration is generally much greater

than on the wing alone because the wetted area is much greater. To make the initial

estimate of subsonic parasite drag AeroDYNAMICTM

used the concept of equivalent

skin – friction drag coefficient, Cfe employing data values from large numbers of

similar aircraft types (see table 4.1 in equation block 3.2(b) above), and the fighter

aircraft value of 0.0035 was used in equations E.15 above to determine CDo.

Total drag was calculated using equation E.12 through calculation of the Oswald‟s

efficiency eo and thereby calculating the values of k1 for both the subsonic and

supersonic cases as described in the pervious section for the range of Mach numbers

considered. From this the drag polar values were calculated as given below in table E-

9 below.

From this data the CD vs CL: Lift over Drag vs CL: and CL ^ 1.5 over CL: values

were generated as shown below in tables E-10 through E-12.



269

Table E-9:- Drag Polar analysis data for the F-24 / A-24.

Mach Cdmin Cdo k1 k2 Mach = 1.5

0.10 0.012505 0.012505 0.16705 -0 CL CD

0.85 0.012505 0.012505 0.16705 -0 1 -0.090175 0.049675

1.10 0.053862 0.053862 0.22418 -0 2 0 0.046832

1.50 0.046832 0.046832 0.34952 0 3 0.157806 0.055536

2.00 0.042685 0.042685 0.49731 0 4 0.315611 0.081648

5 0.473417 0.125168

6 0.631223 0.186095

7 0.67631 0.2067

8 0.631223 0.186095

Table E-10:- CD vs CL analysis data for the F-24 / A-24.

Mach = 1.5

CL CD

1 -0.090175 0.049675

2 0 0.046832

3 0.157806 0.055536

4 0.315611 0.081648

5 0.473417 0.125168

6 0.631223 0.186095

7 0.67631 0.2067

8 0.631223 0.186095



270

Table E-11:- L over D vs CL analysis data for the F-24 / A-24.

M=Mcrit M=1.5 L/D vs. CL

CL L/D L/D Plot Mach No. (L/D)max

1 0 0 0 1 Mcrit 10.83

2 0 0 0 2 1.5 3.87

3 0.157806 9.469333 2.8414841

4 0.315611 10.82903 3.86550806

5 0.473417 9.47878 3.78226356

6 0.631223 7.983598 3.39193757

7 0.67631 7.606415 3.27193701

8 0.631223 7.983598 3.39193757

Table E-12:- CL^1.5 over CL analysis data for the F-24 / A-24.

M=Mcrit M=1.5 CL^1.5/CD vs. CL

CL CL^1.5/CD CL^1.5/CD Plot Mach No. (CL^1.5/CD)max

1 0 0 0 1 Mcrit 6.52

2 0 0 0 2 1.5 2.69

3 0.157806 3.7616711 1.12877308

4 0.315611 6.08367624 2.1716158

5 0.473417 6.52190518 2.60239867

6 0.631223 6.34293187 2.69487892

7 0.67631 6.25537056 2.69077853

8 0.631223 6.34293187 2.69487892

The data generated from the aerodynamic analysis of the common F-24 / A-24

airframe configuration was then employed in the performance: constraint: Vn and

manoeuvre: and weight and stability analysis based on the programmed mission

which was entered through the mission builder interface in the AeroDYNAMICTM

toolset.

This concludes discussion on the aerodynamic analysis of the F-24 / A-24 final

configuration using the AeroDYNAMIC V2.08 toolset.

Documents

Design proposal FB-24 CD Report Final issue version