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Department of Aeronautical Engineering

University of Bristol

UB2008FColossus

Sub-Section 7:

Stability & Control

Authors:

Mike DennisonMartin Bracewell

Becky Hutchinson


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Three-View

General DetailsModel Description High payload, long

range freighter

aircraftList Price Std/HO (2007

$USm)

287.95

Launch 2015

Entry into Service 2020

Accommodtn (space ltd) 6 x PAG, 38 x LD3,33 x PMC118

Accommodtn (payld ltd) 18 x PAG, 4 x LD3,

33 x PMC118

Design CriteriaMax Operating Vmo/Mmo 510 KTAS / 0.89M

Dive VD/MD 528 KTAS / 0.92M

Certified Max Alt. 12530m / 41100ft

Landing Gear VLO/VLE

Max. Flaps VFE 205 KCAS

External Geometr Wing (Canard)Overall Length 79.62m/261.2ft

Overall Height 22.47m/73.72ft

Wingspan 75.0m/246.1ft

(25.5m/83.7ft)

Wing Area (gross) 680m2 / 7319ft2

Wing Area (ESDU) 596m2 / 6410ft2

Canard Area (gross) 111m2 / 1195ft2

Canard Area (ESDU) 71.1m2 / 765.3ft2

Wing/Canard ARatio 9.92 / 5.87

1/4 Chd Swp 33.0 deg (35.0 deg)

t/c - Root / Kink 1 /

Kink 2 / Tip

0.14/ / /0.105

Cabin Geometry

Max no ULDs 14 x PGA or

33 x AMA or

33 x PMC or

33 x AAK

Cabin length

Top:

Upper Cargo:

Lower Cargo:

9.46m / 31.04ft

49.97m / 157.38ft

47.53m / 155.94ft

Cabin volumeTop:

Upper Cargo:

Lower Cargo:

78.30m / 2765ft3 3

1026.2m / 36239.9ft3 3

375.4m3 / 13257.1ft3

Max cabin width

Top:

Upper Cargo:

Lower Cargo:

5.77m / 18.9ft

6.82m / 22.4ft

6.8m / 22.3ft

Cabin floor width

Top:

Upper Cargo:

Lower Cargo:

5.77m / 18.9ft

6.82m / 22.4ft

5.18m / 17.0ft

External

Fuselage width:

Fuselage height:

7.11m / 23.33ft

8.51m / 27.92ft

Total cargo volume 909.3m3 / 32112ft3

Payload net density 58.6kg/m2 / 12 lb/ft2

Colossus

Fuselage Cross-SectionDeck Layouts


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Cargo Doors/Access

Cargo door

size

(Wdth x hght)

upper cargo: 3.4 x 3.18m

11.2 x 10.43 ft

lower fwd: 2.64 x 1.68m

8.66 x 5.51ft

lower aft: 2.64 x 1.68m

8.66 x 5.51ft

SystemsEngine Rolls Royce Trent 900

Or GP7000 Derivative

APU NASA/Boeing Concept 440kW

Solid oxide fuel cell

Avionics Proprietary

Payload-Range Diagram

Payload Range

0

20

40

60

80

100

120

140

160

180

200

0 2000 4000 6000 8000 10000 12000

Range (nm)

Payload(

x1000kg)

Weights & LoadingsMaximum Ramp Weight 557480kg / 1229012lb

Maximum Takeoff Weight 557010kg / 1227976lb

Maximum Landing Weight 473460kg / 1043783lb

Max Zero-Fuel Weight 373960kg / 824427lb

Operationl Weight Empty 198960kg / 438624lb

Maximum Payload 175000kg / 385802lb

Maximum Usable Fuel: 310336L / 81991USG

** 6.75 lb per USG 251040kg / 553439lb

Payload at max. fuel 107480kg / 236949lb

Wing Loading (MTOW) 711.1kg/m2

145.7lb/ft2

Thrust(max SLS)to Weight

(MTOW)

2.49 N/kg

0.254 lbf/lb

Empty Weight/mx payload 1.14

OWE/MTOW Fraction 0.36(MZFW-OWE)/MTOW Fractn 0.31

Max Fuel Fraction 0.45

Performance

Engine Rating 356kN / 80000lbf

Takeoff Rating max 354kN / 79512lbf

Flat Rating ISA+15 at T/O

Airfield Performance (MTOW/MLW)

BFL, ISA+15C, SL 2745m / 9006ft

LFL, ISA, SL 2060m / 6759ftApproach Speed (MLW) 155kts / 0.24M

En route Perf: Climb (AEO, ISA, MTOW br.)

Time to Climb to FL 350 27 min

Time to Climb to ICA 27 min

Initial Cruise Altitude 10670m / 35000ft

En route Performance: Cruise

Long Range Cruise 487kts / 0.85M

High Speed Cruise 510kts / 0.89M

Payload-Range

Reserves Description FAR 121, 200nm range

Design range for given

payload [@ LRC]

5500nm

Block Performance (given payload, ISA, s.a.)Assumptions: Max payload, LRC speed

3000 nm Block fuel 82662kg/182147 lb

Block time 6.58h

TOGW 477730kg/1053197 lb

Max Range Block fuel 158308kg/349000 lb

Block time 11.71h

TOGW 557010kg/1227976 lb


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Systems Description

ATA-21 Air Conditioning

ECS Overview Fully automated

3 ECS packages

4 zones

2 Ram air scoops

ECS Location Belly Fairing

Cockpit / Cabin Pressure

Control

Automatic and manual

Cockpit / Cabin

Temperature Control

Automatic and manual

No. Cabin Control Zones 3

Press. System Overview Digital controller

Fresh Air Ratio 4 recirc fans

Overpress. Valve Diff. 9.1 psi

Cabin Alt. at Max Alt. 8000 ftCooling Cycle Overview 3 ECS packs

4-wheel cyc. Machine

Dual heat exchanger

Water separator

ATA 22 - Auto FlightAuto Flight

Cntrl Descr.

Digital FCCs

Flight Director Descr. 1 FD per FCC

Yaw Damper Descr. Incorporated into

stability

Auto Pitch Trim Descr. Trim via CGmanagement (trim

fuel tanks)ATA 23 - Communications

Comms System Overview VHF radios, HF data

radios, multimode

receiver (MMR), ACPs

ACARS Standard

SELCAL Standard

ATA 24 - Electrical Power

Main Power Type 270V DC

Power Distr. Frequency VariableNumber of Main Genrtors 8

Main Generator Power 200kVA

Aux. Generator & Power

(APU)

200kVA

Emergency Power Source 4 x PMG (100kVA)

Main System DC voltage 270V, 28V

Battery Type & Power Lithium Ion

Number of Batteries 16

Extrnl AC or DC Hook-Up DC

Main Distrbtn System Integrated modular

avionics, full duplex

ATA 25 Cargo Compartments

Cargo Handling System

Overview

Rheinmetall Power Drive

Unit System, central

maintenance computer

Internal Loading

Apparatus

Ball bearings and

powered rollersProvisions for

loading/unloading

Large cargo doors

Strengthening around

doors

Provisions for unusual

freight

None

Floor loading 2823 kg/m / 1897lb/ft

Cargo compartment alt 8000 ft

ATA 27 - Flight Controls

Flight Control System Electrical

Aileron Actuation Mthod Two Section, 2 Dual

power source actuators

er sectionDescription of Rudder Two Section

Rudder Actuation Method 2 Dual power source

actuators per section

Fixed / Var. Incd. Tail Fixed

Elevator Actuation Mthd 2 Dual power source

actuators per section

Stall Protection Devices Envelope protection inFCS

Flap System Overview Three section single

slotted fowler main

wingFlap (Slat) Deflection -

Takeoff (Highest)

Main Wing 20 (40)

Canard 20 (35)

Flap (Slat) Deflection -

Landing Configuration

Main Wing 40 (35)

Canard 40 (40)

HI Lift LE Device 3 panel kruger flaps on

main wing, 1 panel slaton canard

HI Lift LE Dev. Actuatn Electrical

HI Lift TE Device Single slotted

HI Lift TE Dev. Actuatn Electrical

Total Number of Roll

Splers / Flight Splers /

Ground Splers / Total

/ / / 6

Spoiler Actuation Electrical


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ATA 28 - Fuel System

Tot. Usable Fuel Capac. 287290L / 75894 USG

Tank Capacity (Wing) 202420L / 53474 USG

Tank Capacity (centre) 48620L / 12844 USG

Tank Capacity (canard) 36250L / 9576 USG

Tank Cap. (Aux.+Trim) 14110L / 3727 USGFuel System Overview Automated fuel

distribution

Loctn Aux. Fuel Tanks Outer wing

Fuel Pump Overview Automatically pumped

to change weight

distribution

Cross-Feed Capability yes

Single Pt Refuel Capab. yes

Gravity Refuel Capablty yes

Location of Fuel Filler

Ports

below wing leading

edge between engines

ATA 29 - Hydraulic Power

Hydrlic System Overview N/A no hydraulics

Hydraulic Bay Location N/A

Number of Main Systems N/A

Hydraulic Fluid Type(s) N/A

Nominal Working Pressure N/A

Hydraulic Pumps N/A

Hydraulically Actuated

Items

N/A

ATA 30 - Ice and Rain ProtectionAnti-Ice System Overview Electrical, bleed air

for nacelles

Wing Electric heat mats

Canard Electric heat mats

V-tail Electric heat matsNacelle Intake 5th stage bleed air

Probes & Sensors Electric

Windshield Electrically heated

2 wipers for rain

Rain repellent liquid

ATA 32 - Landing Gear

Landing Gear Actuation EMA

Emerg. Extension

Procedure

Manual release, gravity

extensionMain Landing Gear Type Cantilever

Location of MLG Fuselage and Fairing

MLG Strut Type Oleo-Pneumatic

Tire Size - MLG 1.27 m x 0.51 m

50 in x 20 in

Tire Pressure - MLG 14.32 bar / 208 psi

MLG Braking System Carbon brakes, anti

skid system, EHA

Nose Landing Gear Type Cantilever

Spatial Direction for

Retraction of NLG

Forwards

NLG Strut Type Oleo-Pneumatic

Tire Size - NLG 1.27 m x 0.51 m50 in x 20 in

Tire Pressure - NLG 13.65 bar / 198 psi

NLG Steering Overview EHA, controlled by RDCs

ATA 34 - Navigation

No. of ADS Computers 2

Number of AHRS 2 GNADIRS

STD / OPT GPS STD

EFIS Displays Overview Pilot: 4 LAD (15.3in

Number of IRS 3 STD

STD / OPT EGPWS STDSTD / OPT TCAS STD

No. Radio Altimeters 2 STD

STD / OPT HUD STD

STD/OPT CatIIIa Appr. STD

STD/OPT CatIIIb Appr. STD

STD / OPT Autoland STD

GPWS/Wind Shear Detec STD

Digital Weather Radar STD

STD / OPT EVS STD

STD / OPT MLS STD

Number of VHF Radios 3 STDNo. of HF Transceivers 2 STD

No. of ADF Receivers 2 STD

No. DME Transceivers 2 STDSTD/OPT Mode S Trnspn 2 STD

STD / OPT Coupled VNAV STD

RNP Capability 0.3 or better

Overview of FMS System Integrated terrain

guidance and on ground

navigation

ATA 35 - OxygenOxygen System Overview Crew O2 from cylinder,

PAX chemical oxygengenerators


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ATA 36 - Pneumatics

Pneumatic System

Overview

N/A

Location of Bleed Ports

and Capacity

N/A

Pneumatic Source & Use N/A

Bleed Leak Detection N/A

ATA 39 - Electrical / Electronic Panels

Loc. of Major Elec.

Components & System

Main Cabin

Main Display Panels LCD

Main Display Size (HxW) 15.3in diagonal 16:9

GE Aviation LAD

No. Main Display Panels 4 for Pilot, 3 for

Assist

Avionics Suite Designtn TBD

Avionics Suite

Manufacturer

Proprietary

Avionics Rack Location 1 Rack Rear of Cockpit

1 Rack Fore of Wingbox

ATA 49 - Auxiliary Power Unit

Std / Opt APU STD solid oxide fuelcell

APU Designation Hybrid 440kW SOFC

APU Manufacturer NASA/Boeing

APU Location Tailcone

APU Reqrd for Dispatch Yes

APU Operation & Control FADEC

APU Fire Extinguishing Halon

APU Max Start. Altitude 41000 ft

APU Max Oper. Altitude 41000 ft

ATA 52 Cargo Doors

Access for

loading/unloading

Upper and main deck

doors on port side

Lower deck fore and

aft doors on

starboard side

All doors open

outwards

ATA 53, 54, 55 & 57 - Structure

Strctrl Prss. Diffrntl 8.32 psi

Struc. Life cycle/hrs 120000 hrs

Structure Overview Mostly compositestructure

Structure & Material

Fuselage Frame/Floor

CFRP / carbon

sandwich

Struct. & Material -

Nacelle / Pylon

CFRP sandwich /

Titanium


Canard

CFRP wingbox, GLARE

leading edge


Elevator

CFRP sandwich


Vertical tail

Carbon sandwich,

GLARE leading edgeWing Tip Geometry Type Winglet, CFRP

Struct. & Material -Aileron

CFRP sandwich

Struct. & Material - HI

Lift LE Device

GLARE


Lift TE Device

CFRP


Speed Brakes

CFRP

ATA 71-80 - Engine

Engine Manufacturer Rolls-Royce

Engine Designation Trent

Turbofan No. of Stages

Fan/Boost/Compaxial +

Compcent//HPT/IPT/LPT

1/8/6/1/1/5

Number of Engines 4

Mounting Point wings

Max. Takeoff Thrust each

(Std/HO)

356kN/80,000lbf

Flat Rating Temp ISA+15o

Thrust Reversr Overviw Inboard engines only

Bypass Ratio 7.4

Overall Pressure Ratio 39

TSFC at M0.80, FL 350 0.054kg/N/hr,

0.53lb/lbf/hrFADEC or DEEC FADEC

ETOPS Capability N/AExternal Noise, MTOW

(ICAO Annex 16)

Takeoff / Stage 3 Limit 92.4 EPNdB / 106 EPNdB

Sideline / Stage 3 Lim. 89.7 EPNdB / 103 EPNdB

Approach / Stage 3 Lim. 102 EPNdB / 105 EPNdB

Cumltv Margn to Stg 3 29.9 EPNdB

Emissions (ICAO LTO

cycle)

NOx 53.5g/kN

CO 40.9g/kN

Unburnt Hydrocarbons 4.8g/kN


7/30


8/30

Technical Document 07: Stability & Control

Group 4F i

List of AbbreviationsAC Aerodynamic Centre

ARINC Aeronautical Radios, Incorporated

ATA Air Transport Association

CCS Common Core SystemCG Centre of Gravity

CS Certification Specification

EFCS Electronic Flight Control System

EHA Electro-Hydrostatic Actuator

ELAC Elevator/Aileron Control

EMA Electro-Mechanical Actuator

FBW Fly-By-Wire

FCS Flight Control System

FCU Flight Control Unit

FMC Fuel Management Computer

FMGEC Flight Management Guidance and Envelope Computer

LVDT Linear Variable Differential Transducer

MAC Mean Aerodynamic Chord

MCDU Multi-function Display and Control Unit

MTOW Maximum Take-off Weight

MZFW Maximum Zero Fuel Weight

nm Nautical Mile

OEI One Engine Inoperative

RAC Rudder/Aileron Control

SAG Stability Augmentation

SEC Spoiler/Elevator Control

T/O Take-off

WFP Wing-Fuselage-Pods

List of SuperscriptsA Airborne at V

MC

A

G Grounded at VMCG

V Vertical Tail Plane

WFP Wing-Fuselage-Pods


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Group 4F ii

List of Notations

Angle of attack

Local angle of attack

Sideslip angle

CW Sideslip angle induced by cross-wind Control surface deflection

r Rudder deflection

r max Maximum allowable rudder deflection in normal conditions

r CW Maximum allowable rudder deflection to counteract cross-wind

Surfacec Mean chord of control surface

CH Coefficient of hinge moment

CH 0 Coefficient of hinge moment at zero incidence and control surface deflection

CL Coefficient of lift

CL T/o Coefficient of lift at take-offCM Coefficient of pitching moment

CN Coefficient of yawing moment

CT

Coefficient of thrust

CY Coefficient of side force

H Hinge moment

Kr r =

la Mean aerodynamic chord of wing

lV Vertical tail arm

q Dynamic pressureSc Canard reference area

SW Wing reference area

SSurface Area of control surface

VApp Approach speed

VCW Cross-wind speed

VMC Minimum control speed

Vp Control surface volume

xCG Longitudinal position of the CG from the nose

xCG AC Longitudinal position of the CG from the wing AC

yT Lateral distance to outboard engine


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Group 4F iii

Section Breakdown

M. Dennison: 7-1

7-2: All Sections

7-3: All Sections

7-4: All Sections

7-5: All Sections except 7-5-3-1

B. Hutchinson: 7-5-3-1

M. Bracewell: 7-6: All Sections


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Group 4F iv

Contents

7-1 Introduction...............................................................................................................................1

7-2 Longitudinal Stability.................................................................................................................1

7-2-1 Canard Sizing and Positioning.......................................................................................1

7-2-2 Digital Datcom Analysis................................................................................................1

7-2-3 Nose Wheel Reaction....................................................................................................2

7-3 Lateral Stability..........................................................................................................................3

7-3-1 Cross-wind Landing.......................................................................................................4

7-3-2 Directional Stability.......................................................................................................4

7-3-3 Yawing Moment on Take-off Run.................................................................................4

7-3-4 Yawing Moment and Side Force while Airborne...........................................................5

7-3-5 Final Fin Size..................................................................................................................5

7-4 Derivative Empennage Sizing.....................................................................................................6

7-4-1 Canard Sizing for Derivative..........................................................................................6

7-4-2 Fin Sizing for Derivative................................................................................................6

7-5 Primary Control Systems............................................................................................................7

7-5-1 Control Surface Sizing...................................................................................................7

7-5-2 Control Surface Hinge Moments...................................................................................7

7-5-3 Secondary Control Systems..........................................................................................8

7-5-3-1 Trim Fuel..........................................................................................................8

7-5-3-2 Flap Systems....................................................................................................8

7-5-3-3 Slat Systems.....................................................................................................8


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Group 4F v

List of TablesTable 7-1: Canard geometry and structural parameters 5

Table 7-2: Fin sizing parameters 6

Table 7-3: Fin geometry and structural parameters 9

Table 7-4: Primary control surface sizes 11

Table 7-5: Primary control surface hinge moments 11

Table 7-6: EFCS control laws 13

List of FiguresFigure 7-1: Graph of maximum static margin against CG position 3

Figure 7- 4

Figure 7-3: Canard planform 5

Figure 7-4: Tail planform 9

Figure 7-5: Wing actuation system 14

Figure 7-6: Canard and tail actuation system 14

Figure 7-7: Flight control architecture 15


13/30


Group 4F - 1 - M. Dennison

B. Hutchinson

M. Bracewell

7-1 Introduction

[1],

[1]. Applying these definitions to an aircraft shows that, in the design stages, it must beensured that the aircraft will resist undesired changes in attitude and position whilst still allowing

the pilot to implement desired changes. This document outlines the measures taken during design to

ensure that Colossus remains both stable and controllable.

7-2 Longitudinal Stability

Longitudinal stability is the aspect of stability concerned with displacements and attitudes in the x-z

plane, primarily changes in altitude and pitch angle. While both the main wing and canard have

bearing on the longitudinal stability of the aircraft, it is the canard that can most feasibly be

designed to ensure stability, with the main wing being primarily designed as a lifting, rather thancontrolling, surface.

The aircraft has been designed with the intention that the longitudinal stability will meet CSs 25.171,

25.173 and 25.175 of Certification Specifications for Large Aeroplanes [2].

7-2-1 Canard Sizing and Positioning

The horizontal tail sizing technique outlined in the aerodynamics, stability and control section of the

design manual [3] is not suitable for this configuration. This is due to the assumption that during

flight, CG will move backwards as fuel is burnt. However, for a canard configuration, the CG movesforwards as fuel is burnt and so the take-off rotation parameter - usually one of the key sizing

parameters - does not apply. Therefore, a different sizing procedure had to be adopted for the

canard.

Review of literature on canard sizing suggested that, although a canard wing area equivalent to 25%

of the wing area was ideal to minimise drag [4], canards with more than 15% wing area tended to be

overly destabilising. Therefore, an initial canard size of 15% wing area was selected as a compromise

between requirements for low drag and static stability. Further analysis showed that reduction of

the canard area to only 14% of the main wing area produced a reduction in induced drag as well asdecreasing the destabilising influence of the fore-plane.

It had previously been noted that the wake from the canard would have an effect on the lift

distribution of the main wing. To prevent the effect of the wake being too great, the longitudinal

separation of the two lifting surfaces was maximised such that the downwash from the canard

would pass well below the flow field generated by the main wing, thereby limiting or eliminating any

interaction between the two surfaces.


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B. Hutchinson

M. Bracewell

7-2-2 Digital Datcom Analysis

Using the aircraft geometry data, along with aerofoil section and surface planform data, an

approximate model of the aircraft was constructed. The input file is simple and, as such, the

geometry of Colossus had to be vastly simplified in order for the model to be run. While this reduced

the veracity of the model, the data produced is still more accurate than hand calculations as, thesource data being empirical, effects such as viscosity and vorticity are accounted for. The

assumptions made in the construction of the fuselage model were that the constant diameter

section of the fuselage was cylindrical with a diameter equal to the height of the real fuselage and

that the nose and tail cones were not curved, but simple cones. The wings and canard, instead of

having twist and thickness distributions that vary non-linearly along the span, were assumed to be in

two panels, each with linear thickness and twist distributions. Also, features such as the

undercarriage sponsons, engine nacelles and flap tracks could not be modelled, and neither could

the effect of deployed flaps. Although the database did not include data for the aerofoil sections

used for the wing and canard, the inbuilt vortex panel solver could calculate the sectionalcoefficients given the coordinates of the aerofoil surface. The Datcom input file used can be seen in

Appendix A.

A Digital Datcom analysis of this model was then run. From this computation, it was possible to

ascertain the variation of several

importantly from a stability perspective, the values of

CM

and

CL

were calculated at a range of

C

C-LM

produces a graph of static margin variation with angle of

attack. Using this, and varying the input CG position, it is possible to ascertain the foremost and

rearmost CG positions that keep the static margin within a certain range for a given range of flight

conditions.

7-2-3 Static Margin and Allowable CG Range

The static margin range that has been selected is 5% to 35% MAC. The rear limit, 5% MAC, has been

selected as this is the conventional safety margin applied to ensure that the aircraft remains

statically stable. The forward limit has been selected to produce an allowable CG range of

comparable size to competitor aircraft. It was felt during design that one of the crucial parameters

for producing a viable freighter aircraft was that there had to be at least some freedom of cargo

positioning, and therefore cargo CG position. Producing an aircraft with a more restricted static

margin - and therefore centre of gravity position - was felt to have a greater impact on freighter

operations than having a potentially more than desirably stable aircraft. In addition, the large

moment arm to the canard control surfaces and high power of the lifting surfaces mounted there

were felt to ameliorate the risk that the aircraft would be less controllable than comparable aircraft.


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B. Hutchinson

M. Bracewell

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

35 37 39 41 43 45

CG Position (m)

Max.

StaticMarg

in(%MAC)

Figure 7-1: Graph of maximum static margin against CG position

This static margin range corresponds to an allowable CG range of between 39.26m and 41.5m from

the nose of the aircraft. Fig. 7-1 shows how maximum static margin increases as CG moves towards

the nose. If the CG moves further forward than 39.26m from the nose, the maximum static margin

becomes too large, as can be seen in the figure. The result of an increasing static margin is that the

aircraft starts to become overly stable, r ,

eventually preventing crucial manoeuvres such as flaring, thereby presenting a danger to the

aircraft. Conversely, if the CG moves aft more than 41.5m from the nose, the angle of attack at

which the static margin drops below 5% becomes too small and it becomes dangerous to perform

pitch manoeuvres as they may cause the aircraft to become unstable.

Fig. 7-2 displays the variation of static margin with angle of attack for the 5500nm design missionwith 175t payload. As can be seen, the static margin reduces with angle of attack, up to a point, after

which it increases again. This effect is created purely by the flow interaction of the canard wake with

the wing lifting field. As the turbulent flow from the canard impinges on the wing, it decreases the

magnitude of the nose-down pitching moment generated by the wing, reducing the ability of the

aircraft to recover from pitch disturbances. However, above angles of attack greater than 11 at

cruise, the wake passes over the top of the wing, rather than impacting directly on it. This reduces

the effect of the interaction again. It should also be noted, however, that the interaction only causes

a reduction in static margin below the acceptable 5% MAC at around 9 at take-off, 8 at

cruise and over 10 on landing. These angles of attack are greater than those at which the aircraftoperates in any of these mission phases. It can therefore be stated that the aircraft will remain

statically stable at any mission phase, even without the assistance of a stability augmentation

system.


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B. Hutchinson

M. Bracewell

-0.1

-0.05

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

-5 0 5 10 15

StaticMargin(%MAC

)

Take-Off Cruise - First Stage Cruise - Second Stage Cruise - Third Stage

Cruise - End Approach Zero Fuel Landing

Figure 7-2

7-2-4 Nose Wheel Reaction

In addition to the airborne stability calculations, it was necessary to ensure that the reaction load

produced at the nose wheels was acceptable. The load on the nose wheels cannot be allowed to

exceed 15% of the MTOW of the aircraft, as this makes the aircraft both difficult to steer on the

ground and more prone to nose wheel slam or collapse on landing. It also cannot be allowed to be

less than 5% of the MTOW of the aircraft, as this would give insufficient tire friction for effective

steering.

The undercarriage has been designed with these parameters in mind. The two critical cases for the

nose wheel reaction are when the aircraft is at MZFW and at MTOW, as these are when the CG is at

its most forward and aft positions respectively. When the aircraft is at its MTOW, the allowable CG

range is between 40.62m and 44.35m from the nose, relating to 15% and 5% nose wheel reaction

respectively. The CG position when the aircraft is at MTOW, as produced by the weights and balance

team, is 40.785m from the nose [5]. When at MZFW, the allowable CG range is between 37.86m and44.35m from the nose, while the actual CG position is at 39.51m from the nose. As can be seen, the

actual CG positions are both within the allowable CG ranges. It is important to note that the

allowable CG ranges described here do not apply while the aircraft is airborne, only when the full

weight of the aircraft is being supported by the undercarriage. For discussion of the design and sizing

of the undercarriage, see Technical Report 3: Structures [6].


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B. Hutchinson

M. Bracewell

7-2-5 Final Canard Sizing

Details of the canard geometry and structural parameters can be found in Table 7-1 and the

planform of the canard can be seen in Fig. 7-3.

Part Item STD

Canard

(Reference)

Span 25.5m / 83.7ft

Area 111m2 / 1195ft2

Aspect Ratio 5.87

Anhedral 0

Quarter Chord Sweep 35

Taper Ratio 0.3

Mean Aerodynamic Chord 4.76m / 15.62ft

Quarter chord MAC from fuselage nose 12.23m / 40.12ft

Quarter chord MAC height 2.49m / 8.17ft

Lateral MAC location 5.23m / 17.16ftthickness/chord ratio: root, tip 0.14, 0.105

Spar location (% chord, front / rear) 15% / 70%

Fuel Volume (net) 36250L / 9576 USG

Table 7-1: Canard geometry and structural parameters

Figure7-3: Canard Planform


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B. Hutchinson

M. Bracewell

7-3 Lateral Stability

Lateral stability analysis was performed with the purpose of ascertaining the necessary size for the

vertical tail in order to maintain directional stability despite various perturbing influences. Each of

these influences has been considered as an individual sizing case. The fin has been sized to meet all

cases.

The aircraft has been designed with the intention that the lateral stability will meet CSs 25.177 and

25.181 [2].

The following parameters have been used in the calculation of the required fin size. The lateral

derivatives have been assumed to be equal to those of the reference aircraft, the Airbus A330. The

rudder effectiveness factor has also been assumed to be equal to that of the A330, and linear

interpolation has been applied between the data points to give a value of Kr at any rudder

deflection.

Main Wing Area (SW) 680m2

Wing Mean Aerodynamic Chord (la) 10.11m

Span-wise Position of Outboard Engine (yT) 25.70m

Longitudinal Position of rearmost CG relative to Wing AC (xCG AC) -7.65m

Tail Arm (lV) 27.42m

CyWFP

/ -0.43

CyV/ -3.726

CyV

/r 1.892CN

WFP/ -0.462

Minimum Airborne Control Speed (VMCA) 72.02m/s

Minimum Ground Control Speed (VMCG) 72.02m/s

Approach Speed (VAPP) 79.74m/s

Cross-wind Speed (VCW) 12.86m/s

Lift Coefficient at Take-off (CL T/O) 2.604

Thrust Coefficient at VMCA

(CTA) 0.1647

Thrust Coefficient at VMCG (CT

G) 0.1647

Windmill Drag Coefficient at VMCA

(CD WMA

) 0.0013Windmill Drag Coefficient at VMC

G(CD WM

G) 0.0013

0.0873rads

Maximum Rudder Deflectior max) 0.5236rads

Maximum Rudder Deflection in Cross-r CW) 0.3665rads

Rudder Effectiveness Factor at 30 (Kr 30) 0.900

Rudder Effectiveness Factor at 21 (Kr 21) 0.996

Table 7-2: Fin sizing parameters


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B. Hutchinson

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7-3-1 Cross-wind Landing

The fin must be sufficient in size such that the aircraft is capable of maintaining a straight flight path

on landing when subject to a 25 knot (12.86 m/s) cross-wind. The cross-wind induces a side-slip

CW) as calculated below.

)V

V(tan=

App

CW1CW

The re CW is 0.1599. Using this value and the parameters listed in Table 7-2, the

equation below can be used to work out the required fin area ratio such that the aircraft can

maintain a straight approach in a cross-wind.

a

v

r

V

Y

V

Y

CWa

ACCG

WFP

Y

WFP

N

l

l

rr

C

CW

C

l

x

C

C

v

)K+(

)0.25-+(=

S

S

For this case, the required fin area ratio is 0.001335.

7-3-2 Dynamic Stability

The dynamic lateral stability of an aircraft is usually determined by the qualities of the Dutch Roll

mode at low speeds. However, this involves complex analysis involving nine lateral derivatives and

mass distributions throughout the airframe. Therefore, the fin should be sized to provide an

acceptable level of directional, or weathercock, stability. To achieve this acceptable level, whereby

the aircraft recovers from changes in sideslip angle, the value of C should be greater than 1.25 [3].

Using this and the parameters in Table 7-2, the required fin area ratio can be calculated, using the

equation below.

)-25.0+(

C-)0.25-(-=

S

S

a

CG

a

v

V

Y

a

ACCGWFP

Y

WFP

N

l

x

l

l

C

Nl

x

C

C

V

For this case, the required fin area ratio is 0.1548.

7-3-3 Yawing Moment with OEI on Take-off Run

It is a requirement that the fin be capable of reacting the unbalanced yawing moment produced in

the case of one engine failing on the take-off run. This assumes that the undercarriage provides no

side force reaction. The engine that fails produces windmill drag and does not produce thrust to


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B. Hutchinson

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balance the thrust produced by the corresponding engine on the opposite wing. The equation below

calculates the fin area ratio necessary to react this yawing moment.

r

C

l

l

)C+C(

V

K)(=

SS

r

V

Y

a

V

r

alTyG

WMD

G

T

The fin area ratio required for this sizing case is 0.1364.

7-3-4 Yawing Moment and Side Force with OEI while Airborne

When the aircraft is airborne, it is able to both yaw and translate in the lateral plane. Therefore,

equations relating the fin size to yawing moment and side force must be solved together to ensure

that the aircraft reaches a state of equilibrium. Also, unlike the other cases, it is permissible for theaircraft to assume a roll angle, assisting the reaction at the tail fin with a component of the lift force.

In order that the aircraft reach a state of equilibrium, the two equations below must be solved

simultaneously forS

SV ideways component of the unbalanced

thrust and drag to the side force generated by the fin, while the second equation relates the yawing

moments produced by the unbalanced thrust and drag to the reaction moment at the vertical tail.

rr

C

C

C

LV

K+

-C-=

S

S

r

V

Y

V

Y

WFP

Y

)K-(

)C+(C-))25.0-(+(=

S

S

rr

C

C

l

l

l

yAWMD

ATl

x

C

C

V

r

V

Y

V

Y

a

V

a

t

a

ACCGWFP

Y

WFP

N

The result of solving these two equations is that a fin area ratio of 0.1601 is required.

7-3-5 Final Fin Sizing

The largest fin area ratio required by the sizing cases was 0.1601, required to maintain the aircraft in

equilibrium while airborne at VMCA

when one engine failed. When multiplied by the reference wing

area, the reference area of the vertical fin is calculated as 108.86m2.

Although the planform of the A330 tail is provided [3], it was not felt to be suitable for

implementation on Colossus. The relatively large size of the tail meant that if the A330 planform had

been adopted, the root chord of the fin would have been over 15m and the structure at the root


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B. Hutchinson

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could impinged on cargo volume. Because of this, it was decided to adopt a planform with a smaller

root chord. The height was kept constant and the tip chord increased in order to maintain a constant

area. The selected fin planform is described in Table 7-3 and shown in fig. 7-4.

Part Item STD

Vertical Tail

(reference)

Area (including rudder) 108.86

Aspect ratio 1.33

Quarter chord sweep 40.7

Taper ratio 0.5

Mean Aerodynamic Chord 9.5275

Quarter chord MAC from fuselage nose 68.41

Quarter chord MAC height 5.045

Tail arm 24.71

Tail volume 2689.93

Spar location (% chord, front /rear) 15% / 65%

Thickness to chord; root / tip 10% / 10%

Table 7-3: Fin geometry and structural parameters

Figure 7-4: Tail Planform

7-4 Derivative EmpennageConsideration has been taken of the possibility of having to re-size the empennage if the proposed

higher payload derivative is to be put into service. The results of this consideration are detailed here.

7-4-1 Canard Sizing for Derivative

Since the plan for the derivative aircraft, the Colossus-200, is for fuselage plugs to be inserted

forward of the canard and aft of the wing, the distance between the canard and wing will be

identical to that of the original aircraft. This will mean that although the stretched version will

exhibit a reduced range and performance, the canard should not need to be re-sized to maintain


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B. Hutchinson

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stability. The addition of mass in front of and behind the centre of gravity should also mean that the

CG position does not move very far. This should mean that the static margin for the derivative is

similar to the original aircraft. With the increase in fuselage length, it may be possible to increase the

size of the forward trim tank, described in section 7-5-3-1 to provide greater control over the CG

position in flight, further reducing the need for changes to the sizes or positions of the aerodynamicsurfaces.

It is worth noting that, were a lower-payload, shorter-fuselage derivative to be proposed, then the

canard sizing would have to be considered to ensure that the aircraft was both stable and not

susceptible to increased interactions between the wing and canard flow fields. However, no reduced

payload derivative is being proposed, and such consideration is therefore pointless.

7-4-2 Fin Sizing for Derivative

Increasing the length of the fuselage in order to accommodate more payload will increase themagnitude of the WFP derivatives in the lateral stability equations, which would suggest that the fin

will become undersized for the derivative. However, another effect of increasing the length of the

fuselage, specifically between the centre of gravity and the fin, is that the tail arm will increase. This

increase in the tail arm and, consequently, the moment that can be produced by the fin, will

outweigh the effect of the increase in destabilising WFP terms. Therefore, the fin for the Colossus-

200 will not have to be increased in size over the original aircraft fin.

If a reduced payload derivative with a shorter fuselage were to be proposed, the inevitable

reduction in tail arm would most likely necessitate an increase in fin size.

7-5 Control Surface Sizing

It is vital that the control surfaces be sufficient to provide the pilot with enough authority over the

motions of the aircraft and that the actuation systems be sufficient to ensure that the control

surfaces are always capable of achieving the deflections specified by the pilot or flight control

system.

7-5-1 Primary Control Surface Sizing

The control surfaces on Colossus have been sized empirically using the data provided in the design

handbook [3] for the reference aircraft, the Airbus A330. The span-wise position of the surfaces has

been altered slightly, mainly in the case of elevators on the canard which had to be reduced in span

relative to the elevators on the A330 tail. This was due to the need for there to be powerful high-lift

devices on the canard to counteract the pitching moment produced by the wing-mounted high-lift

devices. For more on the design of the high-lift devices, see Technical Report 5: Aerodynamics [7]. It

was decided that the span of the elevators could be reduced, as long as the trailing edge of the

canard remained fully moveable. Since the reduction in span was to make way for flaps, this was

found to be acceptable. The flaps will be deployed at take-off and landing, reducing the lift increase


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B. Hutchinson

M. Bracewell

required from the elevators, allowing the elevator deflection to be reduced. With the elevators not

at their maximum deflection, there is still a reserve of control power for rapid pitch changes, such as

go-around manoeuvres. The position and size of the primary control surfaces is shown in Table 7-4.

The percentages given in the span limit columns refer to percentage semispan.

SurfaceHinge Line Position

on ChordInboard Span Limit Outboard Span Limit Area

Elevator 70%7.87m / 25.82ft

(62%)

12.75m / 41.83ft

(100%)4.238m

2/ 45.62 ft

2

Aileron 76%28.13m / 92.27ft

(75%)

35.63m / 107.04ft

(95%)9.608m

2/ 103.4 ft

2

Rudder 70%0.6025m / 1.98ft

(5%)

11.45m / 37.57ft

(95%)29.40m

2/ 316.5 ft

2

Table 7-4: Primary Control Surface Sizes.

7-5-2 Primary Control Surface Hinge Moments

The control surface hinge moments were calculated using the procedure outlined in the design

manual [3]. The values of CH0,'

CH

and

CH

provided there for the reference aircraft, the A330, were

surfaces using the equations below. In each case, the worst cases of local incidence and control

surface deflection have been taken to ensure that the actuators cannot be overloaded in normal

operation. The results of the calculations, along with the data used to produce them, are shown in

Table 7-5.

pHVqC=H

where:

SurfaceSurfacep

HH

0HH

cS=V

C+'

'

C+C=C

Surface CH0 Deflection Hinge Moment

Elevator -0.0064 -0.0595 -0.30230 -10940Nm / -8069lbf-ft

-15 5280Nm / 3894lbf-ft

Aileron -0.0758 -0.1 -0.586325 -72640Nm / - 53570lbf-ft

-25 41260Nm / 30430lbf-ft

Rudder 0 0.117 -0.43 30 371100Nm / 273700lbf-ft

Table 7-5: Primary Control Surface Hinge Moments


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7-5-3 Secondary Control Systems

Secondary controls are those that are not used for short-term changes to the aircraft position or

attitude. They are instead used to make protracted changes to the aircraft dynamics either by

affecting the CG position in the case of trim tank systems or by significantly affecting the lift and drag

characteristics of lifting surfaces in the case of spoilers, slats and flaps. They usually take longer todeploy or retract, hence not being used for instantaneous or short-term control demands.

7-5-3-1 Trim System

There will be a 36.25 tonne fuel tank built into the canard box to aid aircraft stability. Without this

trim tank, the centre of gravity would drift forwards as fuel in the wing was burned. The solution is

to start with the canard fuel tank full, and as the fuel in the wing is burned, pump the canard fuel

backwards. This will offset the centre of gravity drift.

Fuel flow will be controlled by the Fuel Management Computer (FMC). The FMC will calculate exactly

how much fuel must be pumped back, the rate it should be pumped at and when it should be

pumped. Valves in the fuel flow pipes will be utilised for flow control; the valves themselves will be

controlled by the FMC.

7-5-3-2 High-Lift SystemThe high-lift system implemented on Colossus features high-lift devices on both the canard and main

wing. The leading edge of the canard has plain slats to 95% semispan while the inner 50% of the

trailing edge of the exposed semispan it devoted to double-slotted Fowler flaps. The wing features

Kruger leading edge flaps over 85% of the semispan with breaks for the pylons and single-slotted

Fowler flaps to 70% semispan. For further discussion of the design of the high-lift system, see

Technical Report 5: Aerodynamics (Reference 7) and for the actuation methods employed, see

section 7-6-2 below.

7-6 Flight Control System (FCS)

The flight control section can be categorised into three sections: the architecture, the control laws

and the actuation methods. All three sections are examined here and in more detail in the Systems

Specialist Report.

7-6-1 Electronic Flight Control System (EFCS)

The EFCS is fully integrated into the common core system (CCS) onboard Colossus. The system is a

fail operational design, as there is no manual reversion available for the system. The EFCS controls

the aircraft in roll, pitch and yaw through five cross coupled control processes:

1. Elevator/Aileron Control (ELAC)

This process controls the elevator position and the aileron channel. Control of both

pitch and roll, with secondary control of yaw through asymmetric deployment of the

ailerons and elevators.


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2. Spoiler/Elevator Control (SEC)

Controls the spoiler deployment on the main wing, along with secondary control of

the elevators. This process can control the aircraft in pitch and roll.

3. Rudder/Aileron Control (RAC)

Manages the rudder control channel, along with secondary control of the aileron

actuators.

4. Flight Augmentation Computer (FAC)

Damps unwanted oscillations in the aircraft, such as dutch roll through the yaw

damper and through the primary control surfaces.

5. Stability Augmentation (SAG)

This function monitors the attitude of the aircraft and augments the longitudinal

stability of the aircraft. It has input into the elevators and the fuel management

system to limit the motion of the centre of gravity.

The EFCS implements three different sets of control laws depending on system health.

Normal laws are applied when system health is at 100%.

Alternate law sets are applied if the system performance degrades.

Direct law is used in the event of further system degradation.

This combination of Alternate and Direct law ensures that the EFCS initially fails operational, and

upon compounding failures, fails safe. The details of each law set are shown in table 7-6.

EFCS Control Law Set Services and Limits

Normal Dynamic Envelope protection : pitch limit, roll limit, overspeed

protection

Dynamic Load factor protection: pitch rate, roll rate.

Stall protection

Auto turn co-ordination

Flap/Gear down speed limiting

Dutch roll damping

Phugoid motion damping

Alternate Fixed Envelope protection : max pitch attitude, max roll angle, max

speed

Direct Direct control coupling to control sidestick, no envelope protection.

Table 7-6: EFCS Control Laws


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7-6-2 Flight Control Actuation System

The flight control system is an electrically signalled, electrically powered system, with position

sensing feedback to the EFCS and pilot. The system uses a combination of electromechanical

actuators (EMAs) and electro-hydrostatic actuators (EHAs).

Fig. 7-5 shows the actuation system for the main wing, and fig. 7-6 shows the actuation system for

the canard and tail.

Figure 7-5: Wing actuation system

Figure 7-6: Canard and Tail Actuation System

Flap

Flap

Ailerons

Kruger Flap

Flap

Spoiler

SpoilerSpoiler

E

M

AE

M

A

E

M

AE

M

A

E

M

A EM

A

E

M

A

E

M

A EM

A

E

M

A EM

AE

M

AE

H

A

E

H

A

E

H

A

E

H

A

E

M

AE

M

AE

M

A

E

M

AE

M

AE

M

A

Kruger Flap

Kruger Flap

Primary Power Source

Secondary Power Source

Colour corresponds to power bus, i.e.

red, blue yellow or green 270V DC bus

bar.

TAIL

EHA

EHA

EHA

EHA

Primary Power Source

Secondary Power Source

Colour corresponds to power bus,

i.e. red, blue yellow or green 270V

DC bus bar.

Flap

Slat

CANARD

Elevator

Flap

E

H

AE

H

A

E

H

AE

H

A

E

M

A

E

M

A

E

M

A

E

M

A

E

M

AE

M

A

Rudder

Rudder


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B. Hutchinson

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The primary flight controls (the elevators, ailerons and rudder) are all actuated using EHAs. The

secondary control surfaces and high lift devices (slats, flaps and spoilers) are actuated using EMAs.

Each control surface is actuated by two, dual power source actuators. This provides each actuator

with a quad redundancy, ensuring that if there is power on any main bus bar (probability of nopower on any bus bar 9.45 x 10

-26(Reference: Systems Report, Power Systems) which is the

probability of catastrophic failure of the electrical) one actuator per surface will be powered. This

results in very high system reliability, and allows the reduction in the number of tail and aileron

segments, reducing the number of actuators required and so increasing efficiency.

7-6-2 Flight Control Architecture

The control of Colossus is achieved through two closed control loops. The fly-by-wire flight

control loop, and an autoflight / navigation control loop through the Flight Management

Guidance and Envelope Computer (FMGEC). This control architecture can be seen in figure 7-7.

Figure 7-7: Flight Control Architecture

The fly-by-wire control system (FBW) incorporates the EFCS control laws, and combines the pilot

input demands under manual flight, or the FMGEC commands in autoflight. The FBW computer then

calculates the correct control inputs for the actuators, through the ELAC, SEC, RAC, FAC and SAG

processes.

These control demands are transferred to the actuator controllers via an ARINC 629 data bus. The

control loop is closed via a feedback loop using various sensors, such as Linear Variable Differential

Transducers (LVDTs) and other sensors to determine the actuator position and return the data to the

FBW processes in the CCS via the flight ARINC 629 databus. The sensed position and demanded


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B. Hutchinson

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position are compared and the FBW computer then recalculates a new control signal, taking into

account new demands from the pilot or FMGEC.

The reliability of the system is ensured through a combination of the multiple redundancies within

the FBW Actuator control loop, and the CCS architecture. The FBW computer uses 5 processes whichare all cross coupled, ensuring that the failure of a single process does not compromise the control

loop. Furthermore, the avionics architecture onboard Colossus (Dual flight data busses, Dual CCS)

ensures that the backup architecture will monitor and if necessary replace the primary FBW

computer. This dual redundancy along with the inherent redundancy in the CCS provides a FBW

system that can incur multiple failures without significant performance degradation, effectively the

system will fail operational twice (i.e. each FBW computer on each CCS must experience faults)

before the alternate laws for the EFCS are used.


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References

1: Oxford English Dictionary

http://www.oed.com/

Date last accessed: 22/02/08

2: Certification Specifications for Large Aeroplanes CS25

www.easa.europa.eu/doc/Agency_Mesures/Agency_Decisions/CS-25_Amdt4.pdf

Date last accessed: 22/02/08

3: 2008 Design Handbook Section 7: Aerodynamics, Stability and Control

University of Bristol/Airbus UK

4: Aircraft Design: A conceptual Approach - Fourth Edition

Daniel P. Raymer. AIAA Education Series, New York: 2006

5: Technical Report 4: Weights and Balance

T. Petzold, P. Gallagher and M. Smallwood

6: Technical Report 3: Structures

P. Gallagher, M. Smallwood and T. Petzold

7: Technical Report 5: Aerodynamics

D. Mansell and M. Dennison


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Appendices

A: Digital Datcom Input FileCASEID APPROXIMATE COLOSSUS

$FLTCON NMACH=1.0,MACH(1)=.235,ALT(1)=478,NALPHA=13.0,

ALPHA(1)=-5.,-2.5,0.,2.,4.,6.,8.,10.,12.,14.,16.,18.,20.,$

$OPTINS SREF=680.0,CBARR=10.11,BLREF=75.0,$

$SYNTHS XCG=40.883,ZCG=0.0,XW=7.0,ZW=-3.75,ALIW=3.0,

XH=35.74,ZH=4.75,ALIH=3.0,XV=60.86,ZV=4.75,$

$WGPLNF CHRDR=6.68,CHRDTP=2.00,SSPN=12.75,SSPNE=12.00,

SAVSI=35.0,CHSTAT=0.0,TWISTA=-4.0,DHDADI=0.0,TYPE=1.0,$

$WGSCHR TYPEIN=1.0,NPTS=15.,

XCORD(1)=0.,0.001,0.002,0.01,0.1,0.2,0.3,0.4,0.5,0.6,0.7,0.8,0.9,0.95,1.,

YUPPER(1)=0.,0.0062,0.0086,0.0175,0.0415,0.0510,0.0554,0.0566,

0.0552,0.0513,0.0441,0.0323,0.0154,0.0051,-0.00755,

YLOWER(1)=0.,-0.0062,-0.0086,-0.0175,-0.0415,-0.0511,-0.0556,-0.0564,

-0.0526,-0.0428,-0.0270,-0.0092,-0.0004,-0.0023,-0.00755$

$HTPLNF CHRDTP=3.74,CHRDR=14.4,SSPN=37.5,SSPNE=34.57,

SAVSI=33.0,CHSTAT=0.0,TWISTA=-4.0,DHDADI=-2.0,TYPE=1.0,$

$HTSCHR TYPEIN=1.0,NPTS=15.,

XCORD(1)=0.,0.001,0.002,0.01,0.1,0.2,0.3,0.4,0.5,0.6,0.7,0.8,0.9,0.95,1.,

YUPPER(1)=0.,0.0066,0.00912,0.0186,0.0440,0.0541,0.0587,0.06,0.0586,0.0544,

0.0467,0.0342,0.0163,0.0054,-0.008,

YLOWER(1)=0.,-0.0066,-0.00912,-0.0186,-0.0440,-0.0542,-0.0589,-0.0598,-0.0558,-0.0454,

-0.0286,-0.0097,-0.0004,-0.0024,-0.008$

$VTPLNF CHRDR=12.1,CHRDTP=6.02,SSPNE=12.0,SSPN=12.05,

SAVSI=35.0,CHSTAT=0.0,TWISTA=0.0,TYPE=1.0,$

NACA-V-4-0010

$BODY NX=5.0,

X(1)=0.,8.0,30.0,68.0,78.0,

R(1)=0.,3.55,3.55,3.55,0.,

ZU(1)=0.0,0.0,0.0,0.0,0.0,$

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