NASA Technical Memorandum 74055

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A PRELIMINARY STUDY OF THE PERFORMANCEAND CHARACTERISTICS OF A SUPERSONIC

EXECUTIVEAIRCRAFT

VINCENT R. MASCITTI

SEPTEMBER1977

N/LSANational Aeronautics andSpace Administratton

Langley Research CenterHampton, Virginia 23665

https://ntrs.nasa.gov/search.jsp?R=19780005097 2020-03-13T19:49:19+00:00Z

TABLE OF CONTENTS

FORWARD

SUMMARY

INTRODUCTION

SYMBOLS

CONFIGURATION CONSTRAINTS

CONFIGURATIONS STUDIED

MASS PROPERTIES

STABILITY AND CONTROL

PROPULSION

LOW-SPEED AERODYNAMICS

SUBSONIC/TRANSONIC AERODYNAMICS

SUPERSONIC AERODYNAMICS

MISSION ANALYSIS

TAKEOFF PERFORMANCE

NOISE PREDICTION

CONCLUDING REMARKS

REFERENCES

i

1

2

3

7

8

11

18

21

28

36

38

47

5O

53

59

60

A PREL!MINARY STUDY OF THE PERFORMANCEAND CHARACTERISTICSOF A SUPERSONIC EXECUTIVE AIRCRAFT

Vincent R. MascittiNASA Langley Rese_.rch Ce_ter

SUMMARY

A preliminary design study has been conducted to determine the impact o'Fadvanced supersonic technologies on the performance and characteri_;ticsof a supersonic executive aircraft. Four configurations with differentengine locations and wing/body blending were studied with an advarced non-_fterburning turbojet engine. One configuration incorporated an advancedGeneral Electric variable cycle engine and two-dimensional inlet withinternal ducting. A M 2.2 design Douglas _caled arrow-wing was used through-out this study with Learjet 35 accommodations'(eight passengers).

All four configurations with turbojet engines meet tlle perfon_ance goals oF5926 km (3200 n. mi.) range, 1981 meters (6500 feet) takeoff field length_and 77 meters per second (150 knot) approach speed. The noise levels ofturbojet configurations studied are excessive. However, a turbojet withmechanical suppressor was not studied. The variable cycle engine congigu-ration is deficient in range by 555 km (300 n. mi.) but nearly meets sub-sonic noise rules (FAR 36 1977 edition), if coannular noise relief isassumed. All configurations are in the 33566 to 36287 kg (74,000 to 80,000ibm) takeoff gross weight class when incorporating current titanium manu-facturing technology.

A preferred configuration was not chosen for this study. While the perfor-mance results on the various configurations are encouraging, it is felt tha_significant performance improvement can be established with more effort.However, some uncertainties exist mainly in the prediction of aerodynamiccharacteristics, which can be resolved only by wind tunnel tests through theMach number range. Further detailed system integration studies to the depthdescribed in this status report should await the completion of planned experi-mental programs.

INTRODUCTION

Continuing growth in the general aviation industry may require at somefuture date the need for a small supersonic cruise business jet. Expand-ing international markets for corporations and diplomatic missions forgovernment executives could create a demand for the convenience andavailability that a long range supersonic jet would provide.

In the mid-late 1960's, preliminary designs, reference 1 and 2, werestudied showing that transatlantic ranges could be achieved by Mach 2.0-2.7 aircraft carrying 8-12 passengers, with gross weights in the 33566 to36287 kg (60,000 to 80,000 Ibm) class. The studies included variablesweep or delta-wing technology with afterburning turbojet or turbofanengines typical of first generation SST technology. '!oise characteristicswere not reported. During this time the U.S. was engaged in the SSTnational competition and much of the detailed information on SST configu-rations was either classified by the U.S. Government or company proprie-tary. In 1971, Boeing conducted an unpublished study of a lO-passengerdelta-win_ aircraft, which resulted in 4778 km (2580 n. mi.) range,46947 kg (I03,500 Ibm) gross weight with a takeoff field length of 1615meters (5300 feet).

Soon after the cancellation of the U.S. SST program in 1971, NASA initiateda research program (SCAR) in supersonic cruise technology. In the pastfive years, NASA and the industry, reference 3, have identified solutionsto the performance, economic, and environmental problems associated withfirst generation SST's. For example, arrow-wing or blended delta-wingconfigurations provide transpacific range 8500 km (4600 n. mi.)o Aircraftemploying variable cycle engines with coannular noise relief or low bypassratio turbojet engines with advanced mechanical suppressors would meet the108 EPNdB noise constraints (FAR 1969 edition). Advanced titanium tech-nology such as superplastic forming diffusion bonding would provide substan-tial cost and weights savings compared to the methods available in 1971. Inaddition, an extensive experimental program was conducted by NASA and Douglas,reference 4, on a M 2.2 design in the transonic and supersonic speed range.The results of this experimental program were important to the current studyas will be discussed in the next section.

With a great deal of information available in the open literature, prelimi-

nary design studies of a supersonic business jet were initiated in 1976

conducted by NASA Langley Research Center and Vought Corporation, Hampton

Technical Center. The objective of this study was to determine the impact

of advanced SCAR technologies on the performance and characteristics of a

supersonic executive jet (SSXJET) and to identify technical problem areas.

During the course of this study, four configurations (Mach 2.2 designs)

were studied which incorporated a NASA generated advanced dry turbojet engine

and a fifth configuration using a General Electric variable cycle engine.

Special emphasis was placed upon the evaluation of terminal area noise

characteristics and takeoff performance.

No attempt was made to determine the optimum cruise Mach number for aSSXJ_T class aircraft since that would require an extensive data baseand a cemplete market and economic'study, which was beyond the scope ofthe present effort. Rather, the Mach 2.2 design constraint was chosendue to the availability of high-speed wind tunnel data and models_ anda low-speed experimental data base which will be established in thecoming year.

SYMBOLS

AR: A

ATOTAL

AWE I

b

CD

CD 9

ACDGEAR

CDMI N

CDNE T

CDTEST

CDwAvE

CDo%' CDio0%

_CDpRoPULSION

LCD

CL

CL_

CLg

aspect ratio

total wetted area

component wetted area

wing span

drag coefficient

drag coefficient in ground effect

landing gear drag increments

minimum drag coefficient

net drag coefficient

drag coefficient from test

wave drag coefficient

zero and full suction drag coefficients

propulsion drag increment

zero-lift drag increment

Liftlift coefficient,

q Sre f

lift-curve slope per degree

lift coefficient in ground effect

CLNET

CLTEST

ACL

CM

C.G,

Cu, Cp REF

CREF

DEG

ft

h

h0

it

KA

KL

KS

LID

(L/D)MA X

LI, L2, L_

It

M_, M

4

net lift coefficient

lift coefficient from test

lift coefficient at minimum drag pointon polar relative to CI = O.

pitching moment coefficient,

pitching moment

q Sref Cref

center of gravity

flap knee blowing coefficient

reference mean aerodynamic chord

degree

feet

altitude

wing height for ground effect calcula-tions

horizontal tail incidence anglemeasured

atmospheric factor

Bw

lift oarameter - yph_z_

shape factor

fuselage length

lift-to-drag ratio

maximum lift-to-drag ratio

designate leading-edge devices anddeflections

tail length

freestream Mach number

i_

m

MIL' DEM

NX

Pg

Ph

&P

PSP

q

re..d

SREF

Sht

SeC

t/c

tl, t_., t3, t,

T_._T2, "I"3

W

x

x/c

Xn, Yn _ Zn

meter

mi r_mum demonstrated

number of Mach cutting planes in wavedrag analysis

number of aircraft roll angles used inwave drag analysis

se_ level pre._,sure

pressure at altitude

sor..ic boom overpressure

polar shade parameter

dynamic pressure

radian

wing reference area

exposed, projected horizontal tail area

second

wing thickness ratio

designate model trail ing-.edge flaps anddeflections

designate SSXJET concepts trailing-edgeflaps and deflections

horizontal tail 'volume coefficient,

Sht !t

Sref Cref

aircraft weight

longitudinal coordinate

chordwise fraction

nacelle coordinates

Y_

Ylbl2

cz

_0

OWRP

Y

_e

6FLAP

°D, ai, OL

spanwise coordinate

semispan fractiont

angle of attack of the wing referenceplane

angle of attack for zero lift

wing reference plane angle of attack

Mach number parameter _/M 2= -l

ratio of specific heats, = 1.4

elevator deflection measured from tail

chord plane

flap deflection angle

aircraft roll angle for wave draganalysis

pitching acceleration

Mach angle

ground effect parameters

CONFIGURATIGN CONSTRAINTS

In formulating the study of a supersonic executive aircreft one must considerthe credibility of the data base to be employed. The Math 2°7 dcsigm baseestablished by NASA, Boeing, and Lockheed during the past national sumersonlctransport program was considered for this study, and indeed much of the NASAdata and the technology developed from the SCAT 15F arrow-wing concept has beenused indirectly. However, it was felt that Mach 2.7 was on the high side for asupersonic executive jet since it requires more sophisticated variat_le geometryengines_ inlet and nozzles that would provide low to moderate operating cos Lsfor supersonic transport but larger component development costs. Si_'_ce theeconomics of a SSXJET could be dominated more by initial cost considerationsrather than operating costs, a lower Mach number data base was sought.

An important experimental program (Ref.4) was recently completed for a Mach 2.2transport design. A jointly sponsored NASA/Douglas wind tunnel test programwas conducted at the NASA Ames Research Center at transonic and supersonic speeds.

Results of these tests verified a high level of aerodynamic efficiency for theDouglas design wing (trimmed L/Dmax = 9.1) when incorporated with a transportfuselage carrying 273 passengers. These encouraging high speed results has ledto the generation of a new NASA experimental program at low-speeds, wherein a9.14 meter (30 ft) model of the Douglas concept will be tested in the NASALangley full scale tunnel. The model is currently under construction and isscheduled for testing early in 1978. This model is designed to be of multi-purpose to facilitate fuselage, nacelle and empennage changes that will berequired to test a SSXJET configuration. Because of past and anticipatedexperimental data, the Douglas design wing was chosen for this study. Thescaled geometry, shown in Figure I, has been held constant throughout the studyfor all configurations. This arrow-wing planform offers several advantagesover a delta wing planform similar to that used on the Anglo-French Concordeand USSR TU-144. The planform of the wing essentially determines the drag-due-to-lift as well as the wave drag characteristics of the wing. Drag-due-to-liftis significantly reduced when the leading edge is swept well behind the designMach line and the trailing-edge notch ratio is maximized. Also, because theleading-edge is swept at 1.24 Rad. (71°), a subsonic leading-edge radius can beused to improve low-speed lift and stability requirements. By clipping the wingtip from a pointed arrow the structural span is decreased and eliminates thatportion of the wing where experiment has indicated a breakdown of the theoreti-cally predicted flow. Unsweeping the wing tip in the region of high local upwash,in this case to .99 Rad. (570), improves the subsonic efficiency and pitchingmoment characteristics with little detriment to the supersonic wave drag. Tofurther minimize drag, the camber and twist distribution has been optimized forthe Mach 2.2 cruise condition to minimize draq-due-to-lift. WinQ thicknessratio varies from 2.2 percent at the root to 3.0 percent outboard of thetrailing edge break.

The passenger compartment (eight passengers/two crew) is similar to the Learjet35 shown in Figure 2. While this is not necessarily the optimum payload for asupersonic executive jet, it does represent the minimum passenger size con-sidered feasible for this class of long range supersonic aircraft.

The material chosen for this study is titanium. 6AL-4V titanium alloys at Mach2.2 temperatures are muchmore structurally efficient than the best aluminumalloys, whenused for strength design parts for 70,000 hours design life. Im-proved manufacturing processes such as superplastic forming and diffusion bond-ing and the resulting innovative structural concepts could potentially providepart costs approaching those of aluminum, (Ref. 3).

Performance Goals and GroundRules

The range goal for this study was 5926 km (3200 n. mi.) which corresponds to aNewYork to Paris flight. The prohibition of civil supersonic flight overlandunderscores the requirement for at least transatlantic range. However, manyimportant city pairs such as Los Angeles to Honolulu _t 4074 km (2200 n. mi.)could be serviced by this aircraft at reduced gross weight, lower takeof$power settings, and, consequently, reduced noise levels. The enroute missionand reserve legs used for performance evaluation are shownin Figure 3.

The takeoff field length goal was established at 1981 meters (6500 ft) maximumin order that the aircraft could access a larger numberof airfields. Approachspeed was constrained to 150 knots corresponding to nearly Learjet values.

No noise goals were established initially since noise regulations do not cur-rently exist for civil supersonic aircraft. Also the simultaneous goals of1981meters (6500 ft) takeoff field length, (high thrust/weight) and largerhighly swept wing (low wing loading) would drive the design to relatively lownoise levels in any event. Noise was, therefore, a "fall out" of the perfor-manceevaluation.

With the above configuration constraints and ground rules defined, the problemwas reduced to determining the minimumtakeoff gross weight that will satisfythe design mission requirements.

CONFIGURATIONSTUDIED

The selection of configurations for this study was based on a parametricevaluation of engine type, location, and inlet considerations for a twinengine executive class transport airplane. All versions utilize the scaledwing planform geometry of the Douglas Mach2.2 transport concept. Passengers,crew, baggagerequirements, landing flare angle, takeoff rotation angle, andpilot vision envelope were maintained constant for all study concepts. Twoengine concepts were used, a non-afterburning advanced turbojet engine with amixed compression inlet and a variable cycle turbofan GE21/JI1-BIO enginesupplied by the General Electric Company. Both engines were scaled to therequired thrust level and the corresponding engine weight for each configurationstudied.

8

Feur b,_s;c configurations were evaluated; a clean wing version witl_ aft fuse-iage __oL!nted nacelles (SSXJET) a similar concept with the naceiles mour:tedunc!er the wing (SSXJET I), a blended wing/body versicm with af!. fusc]agenacelJes, (SSXJET il) and an aft fuselage integrated engine installatier_ w':thunder wing Z-dimensional inlets (SS×JET III).

SSXJET

With the wing geometry and passenger compartment fixed, the fuselage geo_;_:i:_,and v_ing location are arranged to provide minimum wave drag at Ma,ch 2o_ a_d tosatisfy aircraft balance requirements. For the initial configuration develog-ment a parametric analysis was conducted for a range of airplane arq_:s weightsand thrust-to-weight ratios with a nominal wing loading of 3830 N/m2 (80 psf)._rom this analysis a selected airplane was subjected to passenger, subsystems,and fuel volume verification.

It was discovered that in order to stay within a reasonable aircraft lengthand equivalent body fineness ratio, the airplane becomes fuel volume limited.This means that when payload and subsystems volumes are accommodated allremaining wing and fuselage volume is used for fuel containment and structure.Figure 4 shows the volume utilization for _he wing and fuselage. This configu-ration has a wing area of 89.65 m2 (965 ftz), a fuselage length of 31.39 m(I03 ft), and a wing span of 12.83 m (42 ft). The general arrangement andinboard profile of this configuration is shown in figures 5 and 6. Two nacellesare mounted on the aft fuselage in a location to provide ground clearance duringlanding flare and takeoff rotation. The aircraft accessary gear box is locatedin the aft fuselage between the engines and is driven by both engines. Becauseof concern with the takeoff field length requirement, this arrangement resultsin a clean wing freeing the entire wing trailing-edge for mechanical flaps.The empennage is a "T" tail arrangement with a variable incidence horizontalstabilizer and geared elevator. Main landing gear is a two strut arrangementwith two wheels per strut and retracts inboard in the fuselage. Nose landinggear is dual wheeled and retracts forward of the crew compartment. The passengercompartment is nominally 1.727 meters (5.667 ft) in diameter and accommodateseight passengers. Baggage space is provided directly aft of the passengercompartment. Flight crew spacing is similar to the Learjet 35 with .192 Rad.(11 o ) pilot vision angle over the nose. Windshield slope was selected tobe slightly greater than the Mach angle at the M2.2 cruise condition. Fueltanks and aircraft subsystems are located to satisfy required center ofgravity location throughout the flight envelope.

The aft fuselage nacelle placement raises questions with regard to adverselocal flow variations, which may vary with angle of attack, high local

Mach numbers at the inlet face, and inlet distortion characteristics leading

to engine unstart at supersonic speeds. High speed flow field measurements

are required to address this problem area.

9

SSXJET I

This configuration is similar to the SSXJET except that the nacelles arelocated under the wing. The general arrangement and inboard profile are shownin Figures 7 and 8, respectively. This engine placement eliminates the localflow concerns mentioned above, since the wing acts to keep the flow alignedwith the inlet through the angle of attack range and reduces the flow velocityto less than free stream values. Lower local Mach numbers result in smallerinlet size and increased pressure recovery, although no credit for thisadvantage was taken in this study. Spanwise and chordwise nacelle location wasselected based on minimum wave drag and structural support considerations dis-cussed in more detail in subsequent sections of this report. Wing and empen-nage areas and geometry remain the same as the previous configuration. Thewing was relocated, however, to satisfy balance requirements and the fuselagecross section area was modified to minimize wave drag at the cruise condition.The "T" tail was retained to reduce the engine exhaust impingement on thehorizontal stabilizer. Because of the thin winu, enqine driven accessoriesare located in the fuselage and are driven by quill shaft from the engine toa right angle gear box in the fuselage.

SSXJET II

For this concept wing/body blending was employed. The advantage of wing/bodyblending coupled with a "wrap around" structure, is low supersonic wave dragby potentially reducing the maximum cross section area and providing moreefficient use of aircraft volume. Figures 9 and 10 show the general arrangementand inboard profile, respectively. Wing and empennage areas and geometry re-mained the same as SSXJET along with nacelle location on the aft fuselage.The wing was positioned higher on the fuselage to facilitate the blending. Aspreviously mentioned, the study aircraft are fuel volume limited for the samefuselage length. Therefore, area added in the wing/body fillet must be removedfrom the fuselage so that the M 2.2 area distribution would remain basicallythe same. Because of the "wrap around" structural requirements for this con-cept, usable fuel volume is slightly reduced relative to SSXJET and SSXJET I,since the airplane is volumetrically efficient. In an attempt to furtherreduce the cross section area of the crew compartment and its effect on super-sonic wave drag, a derivative version of this configuration was developed.This concept, designated SSXJET II T, placed the pilot and co-pilot in a tandemarrangement with the pilot in the forward position. Figures II and 12 depictthe general arrangement and inboard profile. Although a slight improvementin drag was achieved, one major disadvantage in this concept is evident. Dueto the tandem arrangement and canopy requirements, access to the crew compart-ment is limited to external entrance and egress, therefore, isolating thecrew from the passenger compartment.

I0

SSXJET}:11

[ach of the previously discussed concepts used the advancedturbojet engineand were sized to satisfy payload/range and takeoff field lensthrequirements set forth in the 9rQundrules. Takeoff noise ca_cu!ationsfor these configurations resulted in high noise levels for this cla_s ofaircraft as discussed in the noise section. In light of these results: theSSXJ_T Ill configuration was developed in an attempt to satisfy soise cor,--straints. Figures 13 and 14 are the general arrangement and inboard prof_le_respectively. The engine used is the variable cycle GEI2/JIi-BIO engine scaledto the required thrust level. All regualr performance characteristics wereprovided by General Electric together with scaling curves for weight andgeometry. Installed thrust-to-weight ratio was increased from.37 to.51.This installed thrust permits takeoff at derated throttle settings resultingin lower exhaust gas velocity and, consequently, reduced noise levels both withand without coannular suppressive effects. Two dimensional inlets are locatedunder the wing with the engines positioned in the aft fuselage. Again winggeometry was retained but the area was increased to 105 m2 (1130 ft 2) to improvelow speed characteristics and compensate for the reduction in trailing-edgeflap area due to the underwing inlet. Fuselage length was increased .81 meters(2.67 ft) to 32.21 meters (105.67 ft) to compensate for the fuel loss due toengine ducting. Overall equivalent body fineness ratio was maintained as theprevious concepts. A disadvantage of this two dimensional inlet arrangementis possibly boattail and base drag penalties not accounted for in this study.

Configuration Assessment

As previously mentioned, each of these configurations are fuel volume limited.In order to stay within the wing geometry constraints, no modifications orchanges were investigated. A more detailed study would be required to assessthe penalties of wing thickness changes and fuselage integration to provideadditional fuel volume. Also, no attempt was made to determine if the thinwing provides sufficient thickness for control surface actuators and othersubsystem requirements.

MASS PROPERTIES

The objective of the mass properties analysis is to evaluate candidatestructural designs and establish, as realistically as possible, theminimum design gross weight at which a particular configuration canperform the desired mission goal.

II

In this study, the design mission goal was to achieve a configuration havingthe mission capability of carrying a payload consisting of eight passengers,their baggage, and a crew of two; a range of 5926 km (3200 n. mi.) at Mach 2.2cruise speed.

Five configurations of supersonic executive jet transport (SSXJET) seriesaircraft were evaluated. These are: basic SSXJET, SSXJET I, SSXJET II,SSXJET II Tandem, and SSXJET III.

Methods and Criteria

One of the prime requisites for attaining valid design evaluations duringearly conceptual development of an aircraft system, is availability ofaccurate weight and balance data. Obtaining precise mass data would requirea detailed structural weight analysis beyond the scope of this study. However,it is possible, after establishing a sound reference base, to produce firstlevel mass data with reasonably adequate confidence levels. These data, while

not highly detailed, do reveal trends and serve to isolate, identify, andassess impacts resulting from incorporation of variations in design ortechnology.

For this study a Vought Hampton developed computerized mass property estima-ting program, ESBULL, was used in the evaluation of the SSXJET series aircraft.This program is statistically based with empirical modifications. It was de-signed to generate mass properties for multi-engined commercial transportsand was used in previous SCAR studies, references 6 and 12. The weightprediction portion has been correlated with data from Boeing, Lockheed andDouglas methods. A sample of the wing prediction performance for the programversus reported weights is presented in Figure 15. This plot contains severallarge subsonic transports and one small executive transport to correlate witha small scale aircraft. It also includes the data points for three large

supersonic transports from the three companies. A comparison of the VoughtHampton generated weight data compared to data obtained from Boeina, Lockheed,

and Douglas is presented in Table l(a) and l(b) for metric units andengineering units respectively.

The configuration selection, sizing, and mass anal}sis are based on thefollowing mission requirement criteria:

o Payload - 8 passengers with baggage

o Range - nominally, 5926 km (3200 n. mi.)

o Design Cruise Speed - Mach 2.2

o fuel - Conventional Fossil (JP) fuel

12

The structural m_ssanalysis is based o,_:an all titanium primary structure.[_esig_ features el_dconstruction techniques for major compoi_ent_are:

o i_irJg and Aerodynamic Surfaces - Stresskin titanium skin/core sandwichpane_s.

o Fuselage - Titanium skin/stringer/frame construct'!on.

o Landing Gear - Two strut main gear and single strut nose gear strL_ctureor high strength steel alloy.

Engine - Single spool, non-afterburning advanced turbojet with variablegeometry turbine and exhaust nozzle for all except SSXJET II_, whichhas the GE21/JI!-B10 double bypass engine.

The sizing and configuration selection synthesis for SSXJET was performed bygenerating a matrix of candidate aircraft with an array of Design Gross Weights(DGW) ranging from 31751 to 40823 kg (70,000 to 90,000 Ibm) with sea ieve!,standard day, uninstalled thrust-to-weight ratios (t/w) varying from .35 to .55.The resulting data from this program were subjected to mission performanceevaluations. The candidate aircraft _ith the best match of DGW, and thrust-to-weight ratio meeting range and overall aerodynamic performance was selected.

Weights

Variations in mass characteristics due to configuration differences are dis-cussed on an individual basis using basic SSXJET for comparison. Each configu-ration is described briefly in order to identify salient design features whicheffect mass characteristics.

The design mass characteristics and the group weights of the selected configu-rations are summarized in Tables II and III, respectively.

Basic SSXJET

As described in Configuration Studies portion of this report, the basic SSXJETis an arrow-wing, eight passenger airplane with two aft fuselage mounted dryturbojets. SSXJET has a maximum design takeoff gross weight of 35720.4 kg(73750 Ibm).

SSXJET I

The major difference between basic SSXJET and SSXJET I is the engine location.On SSXJET I the engines (same engine as on SSXJET) are mounted beneath the wing,whereas on the basic SSXJET the engines are mounted on the aft fuselage.

13

This configuration change results in lower structural weights for both wing andfuselage for SSXJETI. The wing structural weight decreased because the enginepods contributed additional wing bending relief. Fuselage weight was decreasedbecause removal of engines from the fuselage reduced aft fuselage structuralloads. The combinedeffect of decreases in structural weight and recyclingresulted in a design gross weight of 34926.6 kg (77000 Ibm) for the SSXJETI.This is approximately 807.4 kg (1780 Ibm) less than the basic SS×jET.

SSXJETII and II Tandem

The SSXJETII versions are nearly identical to the Basic SSXJETin overallconfiguration geometry, but differ in structural design. Both SSXJETIIversions are blended wing/body concepts, which also differ from one another.

Oneversion, designated SSXJETII Tandem,has a tandemcockpit arrangementwhere the pilots are seated one behind the other, similar to certain jetfight/trainer aircraft. The other version, designated SSXJETII, has a con-ventional side-by-side cockpit arrangement, characteristic of transportaircraft.

The blended wing/body concepts exhibit lower weights than the Basic SSXJET.The major difference occurs in the wing structural weight and is attributableto three factors: (I) blended wing/body design concepts result in relativelydeep wing-root sections, resulting in high thickness to chord ratios andreduced wing weight; (2) reduction in the design gross weight (DGW)resultingfrom the decrease in fuel capacity 1270 kg (2800 Ibm) less than Basic SSXJET(3) the downwardcascading effect on DGWduring recycling for sizing.

The combinedeffect of these factors, which translates into deeper, more effi-cient wing-root structure, lower wing loadings, and smaller engines (the wingarea and t/w were held constant) resulted in a design gross weight of 33565.9kg (74000 Ibm) for the SSXJETII. This is approximately 2154.6 kg {4750 Ibm)less than that of the Basic SSXJET.Weight differences between the SSXJETIIversions due to forebody shape variations are too subtle for detection by thefirst level statistical methods. Therefore, only one set of weight values isshownto represent both the SSXJETII and SSXJETII Tandemconcepts.

SSXJETIII

Three distinct configuration changesdistinguish the SSXJETIII from the BasicSSXJET. These changesare: (I) increased wing area from 89.65 to 104.89 m_(965 to 1130 ft2 ); (2) GE21/JII-BIO double bypass engines in lieu of scaledadvanceddry turbojets; and (3) two-dimensional air intake and long-duct airinduction in lieu of axisymmetric short-duct nacelle system.

14

Theseconfiguration changes result in a higher net structural weight forSSXJETIII. Although wing area increased, wing structural weight decreascdbecauseof two reasons: (I) additional wing structural bending relief wasgained from a 28 percent increase in winq fuel, and (2) wing loading wasreduced from 3878 to 3352 N/m2 (81 to 70 Ibs/ft2). Propulsion _roup weightincreased due to installation of the more complex GE21/JII-BIO double bypassengines which are both larger and heavier than the scaled dry turbojets.The most significant weight increase occurs in the nacelle group. By designand configuration, the two-dimensional air intakes are located beneath thewing about mid-body, while the engines are located at the extreme aft fuse-lage. The result is that the air induction ducts leading from the inlets tothe engine are unusually long and therefore heavy. The combined effects ofall changes results in a design gross takeoff weight of 36081 kg (79545 Ibm)for SSXJ_F III, which is 360.6 kg (795 "Ibm) greater than that of the BasicSSXJET.

Balance

Mass balance characteristics are a major parameter influencing design, configu-ration development, and flying qualities of all aircraft. Because of thebroad operational requirements of supersonic aircraft, (flight at subsonic,transonic and supersonic speeds), the trim and stability requirements are moreextensive than those for subsonic applications.

In order to attain the desired longitudinal stability characteristics andflying qualities in the design of fixed wing aircraft, it is desirable toposition the wing in such a manner that the aircraft center of gravity islocated as close to the aerodynamic center-of-lift as possible. Under flightconditions, the locations of the aerodynamic center migrates as a functionof Mach number while the location of the airplane C.G. remains fixed. Insubsonic applications the variation between the two centers is small, butin supersonic operation the aerodynamic center-of-lift migrates aft as the Machnumber increases causing the longitudinal stability to increase to undesirablelevels. On conventional subsonic aircraft, this unbalance of forces actingthrough the two centers can be trimmed (reduced) by deflection of controlsurfaces such as elevators, trim tabs, canards, or all-moving tail-planes.Deflection of surfaces for purposes of trimming creates trim drag, which ifvery large, is highly undesirable for a long range aircraft. These dragpenalties effect supersonic aircraft performance to a greater degree than thatof subsonic aircraft and a more efficient way of achieving aerodynamic trimmingmust be employed.

The arrow-wing geometry developed by NASA after extensive wind tunnel testing,employs wing camber and twist. The gentle camber slopes used in place oflarge control surface deflections reduce trim drag significantly. The amountof reduction achieved in this manner, by twist and camber alone, is still notsufficient if the C.G. were to remain in the same location for both subsonic

and supersonic flight since some fairly high static margins would still needto be trimmed.

15

A fuel managementsystem (fuel transfer system) is chosen as a meansto augmentthe benefits of wing twist and camber. By using this system to pumpfuel intoaft tanks, as the Machnumber increases, the C.G. can be movedaft to coincidewith the aftward migration of the aerodynamic center and minimize trim unbal-ance. A system of this type, utilizing fuel as a form of useful ballast, isused by such current supersonic transports as the "Concorde" and Tupolev 144,am_m111tary aircraft such as the B-l, SR71,and YF12. This system is inoperational use and is a proven methodof achieving C.G. control.

In each of the SSXJETconfigurations, the wing is located in a mannerwhichtailors the balance characteristics to provide the center of gravity limitsrequired for stability and control as specified in the Stability and Controlsection.

The balance characteristics of each configuration are tailored so that boththe aft C.G. location desirable for rotation/flare characteristics at takeoffand landing, and the more forward C.G. location providing minimumdrag atcruise are attainable.

Combinations of fuel tank sequencing were investigated to determine the mostforward and aft center of gravity excursions and define limit boundaries. Theattainable C.G. excursion envelope with the desired cruise C.G. path providingminimumdrag during a typical mission for each configuration are presented inFigures 16, 17, 18, and 19. All of the points along the desired flight pathlie within the limit boundaries and are attainable by proper fuel management.

All the preceding commentsapply in general to all of the configurations but afew commentsare necessary to point out somedifferences betweeneach of theconfigurations.

Basic SSXJET

Wingapex located at F.S. 226.8, aft fuselage mountedengines.

SSXJETI

Wing relocated with apex at F.S. 206.5 to compensatefor engines beingrelocated from aft fuselage to beneath wing mountedposition.

SSXJETII and II Tandem

Wingapex located at F.S. 226.8, sameas on BasicSSXJET, balance charac-teristics vary only slightly between versions due to minor differences inmassdistribution.

SSXJETIII

Dueto configuration and engine changes, wing apex is located at F.S. 220.SSXJETIII marginally satisfies C.G. limits required by stability andControl.

16

Inertia

The inertia characteristics of an aircraft, which are closely related _.o thebalance characteristics by being a function of the mass distribvtion, areequally important in determining aircraft configuration and flying quaiities.The magnitude and type of inertia forces and accelerations which rnust be over--come and controllable by the aerodynamic control surfaces during flightm_neuvering determines the size of the control surfaces. Response rat.'-'., ameasure of flying quality which is the rate at which the aircraft undtrgoe_ _:change in altitude and direction about its axes within a given time peried,is also governed by the magnitude of the inertial forces and in turn affectssize and capacity of the flight control system_.

SSXJET configurations do not differ greatly in configuration geemetry fromexecutive/business jets currently in operation. Several subsonic jets haveaft fuselage mounted engines. As explained in the balance section, the super-sonic aircraft configuration is more sensitive to balance and inertia consider-ations. The fineness ratio and area distribution required for Mach 2.2 cruisespeed requires a long fuselage. Mounting of the engines on the aft fuselageat the extremity of the airframe resul_s in a type of "dumbbell" distribution,with large discrete masses located far from the aircraft C.G. resulting inhigher pitch and yaw inertias. Differences in inertias and principal axes,which are due to variations in geometry and mass occur among the configurations.Using Basic SSXJET for comparison, some of the more pronounced differences maybe explained by reviewing each configuration individually.

SSXJET I

Inertia values differ from basic SSXJET due to redistribution of mass resultingfrom relocating engines from aft fuselage mounted to underwing oosition.The roll inertia from SSXJET I is higher than Basic SSXJET because the enginepods are farther outboard at the underwing location. However, pitch and yawinertias are lower because moving engines closer to aircraft C.G. reduced thetypical "dumbbell" distribution characteristic resulting from locating largediscrete masses at the extremities of the airframe.

SSXJET II and II Tandem

Inertia values differ only slightly from Basic SSXJET due to reduced weight andminor variations in mass distribution.

17

SSXJETIII

Inertias increase over Basic SSXJETdue to increases in gross weight andoverall dimensions, and the resulting redistribution of mass. The roll andyaw inertias are higher as a result of increased wing span and a greateramountof fuel located in the wings. Heavier engines located at the aftextreme fuselage, and dimensional increases in wing span and overall lengthcontribute to increases in pitch and yaw inertias.

The inertia data for takeoff and normal landing conditions are summarizedinTable IV.

Summary

The use of stresskin in the weight estimation was deemeda conservative approachbecauseof the existance of other advancedmaterials and construction techniqueswhich promise equal or lighter weights. Of the alternate materials and tech-niques under consideration, one in particular, seemsextremely promising.This is a new and innovative patented process, developed by Rockwell Interna-tional, which combinessuperplastic forming and diffusion bonding into a singleprocess (Reference 34). This process, knownas SPF/DB,capitalizes on a uniqueproperty of titianium which permits very large tensile elongations under properconditions of temperature and strain rate. Diffusion bonding is the joining oftitanium under pressure at elevated temperatures without melting or use of bond-ing agents. Through a natural phenomenonsuperplastic forming and diffusionbonding can be accomplished under identical parametric conditions. This is thebasis for the SPF/DBprocess. This process holds greater promise because itreduces componentmanufacturing costs by decreasing part count and labor, andresults in weight savings due to inherent structural efficiency. It isestimated that a i0 percent reduction in airframe structural weight of theBasic SEXJETcould be attainable through application of the advanced technologySPF/DBprocess.

STABILITYANDCONTROL

The horizontal and vertical tails of the SSXJETwere originally sized by usingtail volume coefficients based on the previously studied ASTseries of super-sonic aircraft (Reference 6). This section presents the analysis which wasdone in order to confirm the validity of the original estimate of horizontaltail size and to determine the operating center-of-gravity limits.

The horizontal tails of supersonic cruise configurations have traditionallybeen sized by the takeoff and approach control requirements. A similar pro-cedure has been used for this study. The aerodynamic data used is based onwind tunnel data with theoretical corrections applied in order to account for

18

differences between the tested configuration and the SSXJET. The horizontaltail was assumedto be all movable with a geared elevator. A flap deflectionof 30 degrees for both takeoff and approach was used for the stability andcontrol analysis in order to be compatible with the takeoff noise character-istics presented in this study.

Cri tempia

The analysis was based on the following criteria:

Takeoff

Takeoff center-of-gravity range of 0.544 m (21.6 inches) which

is equivalent to O.06D Cref-

o Takeoff forward center-of-gravity limit determined at the posi-tion for neutral stability with the landing gear located suchthat nose wheel lift-off can be achieved at a speed of 162 kts.

A]_proach

o Approach speed defined at a lift coefficient of 0.6 for a flapdeflection of 30 degrees.

o Minimum demonstrated speed defined at a 0.5 g incrementalmaneuver from trim at the approach speed.

Aft center-of-gravity limit based on the abilitx to provide anose-down pitching acceleration of 0.24 raa/sec L at the minimumdemonstrated speed and the maximum landing weight.

Analysis and Results

The basic data for this analysis were obtained from reference 4. Lift andpitching moment increments due to leading and trailing edge flap deflectionswere obtained from reference 10 and corrected for flap area by the method ofreference 12, Using the data for tail-on and tail-off runs from reference10, a zero incidence tail contribution to lift and pitching moment was obtainedas a function of angle of attack. The horizontal tail control power data were

constructed by using the data of reference 6 for a geared elevator type hori-zontal tail with areas of 4.18 m2 (45 ft ) and 6.04 m (65 ftL), The dataindicated that this type horizontal tail with a tail incidence of _0 degreesand an elevator incidence of -+25 degrees deflection would give maximum controleffectiveness. The resulting low-speed lift and pitching moment data are shownin Figure 20 for a flap deflection of 30 degrees and maximum control deflections.

19

The takeoff center of gravity limits were estimated on the basis of the stabil-ity and control criteria noted previously, the takeoff gross weight and pitchmomentof inertia, takeoff CLMAXin ground effect for takeoff and out of ground

effect for climb out. The approach center of gravity limits were based on theapproach stability and control criteria at the normal landing weight. Applyingthe approach criteria of trimming a nose-up pitching acceleration of 0.24 rad/secat the minimumdemonstrated speed results in an aft center of gravity limit of0.5350 CREFfor a horizontal tail with an area of 6.04 m2 (65 ft z) and 0.4846 CREF

for 4.18 m2 (45 ft2). The landing forward center of gravity limits were deter-mined by establishing the maximumnose-up trim capability at the minimumdemon-strated speed. These forward limit_ were 0^3384 CREFfor the 6.04 mz (65 ft _)tail and 0.3805 CREFfor the 4.18 mL (45 ft L) tail. The neutral stability centerof grayity locations were determined to be 0.4695 CREFand 0.4320 Cm_Ffor the6.04 mz (65 ft 2) and 4.18 m2 (45 ft 2) horizontal talls,respectivelyV_-Thesecomputedcenter of gravity locations are plotted in Figure 21 as a function ofhorizontal tail volume coefficient. Using the criteria that the takeoff forwardcenter of gravity location is 0.549 m2 (21.6 inches) ahead of the aft center ofgravity limit and is at the position for neutr_l stability, a horizontal tailvolume coefficient of 0.1128 or area of 5.76 m_ (62 ft 2) is required. Thecorresponding limits are 0.5216 CREFfor'the aft center of gravity, 0.4615 _REFfor the takeoff forward center of gravity, and 0.3440 _REFfor the approachforward center of gravity. This aft limit is more forward than the aft limitscharacteristic of the ASTconfigurations becauseof the increased pitch raterequirement during approach for the lighter SSXJET.

The SSXJETis statically unstable over the greater portion of its center ofgravity range and the data indicates that the basi_ airframe would have consid-erable pitch-up tendencies. Therefore, a hardened stability augmentationsystem (HSAS)would be required to stabilize the aircraft and achieve acceptablehandling qualities. The term "hardened" means that the stability augmentationsystem has sufficient redundancy to preclude loss of the system. Thesesystems, although not used very frequently in the past, are being used onmilitary aircraft and maysoon be commonon commercial aircraft.

The stability and t_im curves were developed for a projected exposedhorizontaltail area of 5.76 m (62 ft _) and are presented in Figures 22 and 23. FromFigure 22 it can be seen that there is adequate control to produce a nose downpitching acceleration of 0.24 rad/sec at the aft center of gravity limit.The approach forward center of gravity limit has been changedto 0.3611CREFin order to be able to trim the aircraft during approach at a 6 degree angleof attack. This was considered to be the maximumangle of attack for whichadequate visibility over the nose could be achieved. From Figure 23, it can beseen that adequate control exists to trim in the takeoff and climb out configu-ration.

2O

Summary

A horizontal tail with an area of 5.76 m2 (62 ft 2) would provide the SSXJETJ _ "

configuration with sufFicient control during ta_eof. , cllmb-.out and approach.This compares favorably with the 6.04 m? (65 ft-) area originally estimated.

The center.-of-gravity range for this configuration is from 36.11 to 52.16percent of the reference mean aerodynamic chord. The data indicate that theSSXJET configuration would have pitch up tendencies and would require ahardened stability augmentation system in order to have acceptable handiin_qualities.

PROPULSION

To conduct airplane performance and acoustic studies, engine performance datasufficient to encompass the airplane flight envelope must be provided. Listedbelow is the minimum engine performance data required to properly conductairplane performance analysis:

1. Mission performance - gross thrust, ram drag, and fuel flow at variousaltitudes amd Mech numbers.

A. For climb performance data points for two different altitudesat each maximum cruise power levels.

B. For cruise performance - data points for five Mach numbers andaltitudes at maximum and four-part power-cruise settings.

Takeoff, landing, and acoustic performance - net engine thrust,exhaust gas flow rate, exhaust gas jet velocity, exhaust gas exitarea: and exhaust gas exit temperature. These data must be providedfor all combinations of three altitudes, three Mach numbers: at take-off power and four-.part po_er settings.

Engine performance data requirements are shown graphically in Figure 24. Inthis figure the engine data requirement is superimposed on a plot of the Mach-_Ititude climb schedule used in the SSXJET airplane studies.

installed engine performance data for an advanced NASA turbojet engine, asdescribed above, was provided for each SSXJET configuratior studied. Engineperformance data provided by the General Electric Company for the GE21/JI1-BIOengine was used in the study of the SSXJET III airplane. The data repomted inthis document, however, are advanced turbojet engine performance used on theSSXJET configuration at a standard +8oc day and GE21/JII-BIO engine performanceat standard day conditions for missions analysis and for both engines at stan-dard +I0oc conditions for takeoff, landing, and acoustic analysis. No datawere available for the GE21/J11-B]O engine at standard +8oc conditions.

21

NASATurbojet Discription

The advancedturbojet cycle employed in the SSXJETstudies is one which wasgenerated at the NASALangley Research Center, as an improvementover theGE4/J5Pengine which was designed for the United States Supersonic Transportporgram.

Turbojet engines are the most fuel efficient engines available for supersoniccruise conditions, however, are less efficient at subsonic cruise conditionsand more noisy than turbofan engines at takeoff. It would be desirable, there-fore, to have an engine which operates like a turbofan at takeoff and subsoniccruise condition; for example,a variable cycle engine.

Howeveran attractive turbojet cycle can be conceived using a variable geometryturbine, and thus approach the variable cycle capability. The turbojet cycleused in this study is discussed in somedetail in reference 5.

Basically, the variable geometry turbine turbojet permits the engine to operateat the design compressor pressure ratio and efficiency at a reduced turbineinlet temperature (reduced power setting) by adjusting the turbine inlet area.With the engine operating at these conditions, at part power, the engine opera-ting efficiency is higher than it would be if the turbine inlet geometry werefixed. This results in a lower fuel consumption at part-power operating con-ditions encountered during subsonic cruise.

Another feature of the variable geometry turbine turbojet is that while opera-ting at part power, the engine airflow is higher than with a fixed turbineturbojet. This is a desirable feature for jet noise reduction. Jet noiseproduction is simply related to the engine exhaust gas velocity. Engine thrustis also related to both the exhaust gas velocity and flow rate. Thus, whenthe gas flow rate is increased at the samethrust level, the exhaust gas velo-city is reduced with a resulting reduction in jet noise.

This engine cycle has been used in several supersonic transport design studies,one of which is reference 6.

A scaled version of this advancedturbojet engine has been used in the SSXJETstudies and has the following sea level static standard day design charcteris-tics:

Overall compression ratio = 15:1

Turbine inlet temperature = 1700OK(3060°R)

Corrected engine inlet airflow = 72 kg/sec (158.7 Ibm/sec)

Uninstalled gross thrust = 73000 N (16410 Ibf)

22

Bare engine mass (weight) = 881 kg (1942 Ibm)

Th_s bare engine massincludes the gas generator, nozzle, and thrust_'ew;r'ser.

._ s_:etc/_of tilis engine is shownon Figure 25.

The r_elected engine is not necessarily the optira_uff: engine for a Mach 2.2 c'.-uiseSSX,_FT a_rp:an,':. As --hown in reference 5 it may be possible to select ._Jnengine with better performance by choosing one with a higher" overall con:prGssorpress_me ratio. To determine the optimum engine for a particular airplane mis-sion wou_d require a complete evaluation of the entire desired airplane flightenve!ope for eacl; availabie engine installation. Such an engine/airplane opti-mization was not accornplished in this study. Optimization, however, shc_u!d beundert_Lnn in future studies.

A r,aceTIe to house the enq'ine was configured as shown fn Figure 26. The resuitsof an unpublished study which evaluated the wave drag characteristics of variousshape nacelles indicated that the configuration selected would have the minimumwave d_-ac,,

The cylindrical nozzle exit would also result in no boattail draq. In order tohave a cylindrical nozzle exterior.and still have a converging diverging fullyexpanded nozzle, an ejector nozzle was utilized. An ejector nozzle is one whichinduces a secondary flow into the nozzle. The secondary flow not only cools thenozzle waiT_ but fi_Ts the void between the nozzle walls and exhaust gas stream,thus, minimizing nozzle losses. The large flight envelope of the SSXJET airplaneis such that it is not possible to maintain a cylindrical nozzle exit at allflight conditions. An ejector nozzle does, however, minimize the nozzle boat-tail angle at those flight conditions where a nozzle boattail is necessary.

The inlet used is a scaled version of the NASA Ames developed "P" inlet ofreference 7. This is an axisymmetric inlet with a translating spike. Thet_'ansl_ting spike is provided to match the required engine airflow at allflight operating conditions with minimum losses.

The maximum diameter of the nacelle was determined by the fully expanded nozzleexit diameter at the design cruise conditions of:

Mach number : 2.2

Altitude = 18288 m (60000 ft)

Atmosphere = Standard day

Other nacelle dimensions are based on the AST-].O0 nacelle and the GE4 engine ofreference 6.

As the SS×JET studies progressed, the size and configuration of the airplanechanged resulting in different size engine requirements. Different engine and

23

nacelle sizes were determined by means of the scaling data shown in Figure 27

from reference 9.

Installed performance of the engine was generated with the aid of the computerprogram of reference 9. The computer program generates engine performance by

means of a cycle match procedure. It employs the GE4/J5P compressor map in anon-dimensional form, fixed combustor efficiency, and a fixed turbine polytro-

pic efficiency. The design point calculation fixes the compressor design point,

dimensionalizes the compressor and establishes the matching turbine flow func-

tion. The turbine flow function remains fixed and an engine operating point is

determined by flow matching the compressor and turbine at a given turbine inlet

temperature and engine rotor speed.

Installation effects of the inlet, nozzle, thrust reverser, service air bleed

of 0.454 kg/sec (i Ibm/sec), and 149 kw (200 HP) power extraction are includedin the calculation of installed engine performance.

Drags due to inlet spillage, bypass, nozzle boattail, and air conditioninm areaccounted for by applying a drag increment to the airplane drag polars. ?he

combined drag increment for spillage, bypass, and nozzle boattail were esti-

mated by scaling the propulsion drag increment from reference 6 based on refer-

ence wing area and engine airflow. The total drag increment is shown on Figure28.

The installed performance data for the advanced turbojet at standard +8oc

atmosphericconditions is presented on Figures 29 throught 31. Superimposed on

Figure 31 are the beginning and end cruise operating points. Data for takeoff,

landing, and noise studies at standard +I0oC atmospheric conditions is provided

on Figures 32 through 35.

Design sea level static thrust and airplane thrust to weight ratios at various

conditions are tabulated below for the engine and the SSXJET airplane at a

design gross weight of 35720 kg (78750 Ibm).

Conditions Thrust Thrust/Weight

Uninstalled standard dayInstalled standard daYInstalled standard +8°C dayInstalled standard +lO°C day

72995 N (61410 I bf)67577 N (15192 Ibf)64677 N (14540 Ibf)64000 N (14388 Ibf)

.4168

.3858

.3693

.3654

General Electric Variable Cycle Engine Description

The GE21/J11-BIO engine is a double bypass variable cycle turbofan engine based

on technology available in 1985. Double bypass variable cycle describes the

function of the engine which, for takeoff and low-speed operation increases the

bypass airflow to provide better low-speed performance both acoustically and

thermodynamically. For high-speed performance, the bypass air is reduced to

provide performance approaching that of a turbojet engine.

24

An addition_l feature of this engine is that the bypass airflow, which inother turbt_fan engines would be tile outer stream, in this engine is divertedso that it e×hausts through the nozzle plug and becomes the inner stream. Thisis done to gain the effects of coani_ular sound suppression (reference 8),

Geometric and engine performa_c.£ data for the GE21/JI1-B10 engine were providedby the General E_ectric £ompa:qy. This engine is designed for a crL_ise Machr.umber of 2,2 at an altitude of 18288 m (60000 ft) for standard atmasphe>'ic con,-ditions with the following design characteristics:

Bypas_ ratio = 0.35

Overall pressure ratio = 17,3

Maximum afterburnin9 temperature = 1355°K (2439°R)

Corrected engir, e inlet airflow = 317.5/349.3 kg/sec (700/770 Ibm/sec)

Engine mass (weight)

Gas generator, afterburner, annular

Nozzle and suppressor 5343 kg (11780 Ibm)

Afterburner system 118 kg ( 260 Ibm)

A sketch of the GE21/J1]-B10 engine scaled to the SSXJET III size is shown onFigure 36.

The engine as provided by GE is much too large for the SSXJET ]iI airplane;therefore, it was scaled to the required uninstalled thrust of 96526 N (21700ibf) at sea level static standard day conditiors. Scaling was accomplished bymeans of the scale factors suggested by GE as shown below:

Weight

!Wa2 _I. IWt : Wt1

"2 ,Wa] ]

Diameter

= D1Fn2 I" 5

].ength

Wa i "5I__ =; LI _._-I

25

Where

Wt = Engine mass (weight)

D = Diameter

L = Length

Wa = Engine airflow

Fn = Net engine thrust

Subscript 1 - GE21/J11-BIO parameter

Subscript 2 = Desired engine parameter

Mass (weight) of the scaled engine is 1828 kg (4029 Ibm) and includes the gasgenerator, afterburner, and annular nozzle with a built in thrust reverser.

In general, scaling an engine more than 25 percent either up or down is

inadvisable. The above scaling factors provided by GE, ho_vever, are

considered sufficiently accurate for the purpose of this study.

The performance data supplied includes the installation effects of inlet, nozzle,

thrust reverser, afterbody drag, service airbleed of 0.454 kg/sec (l Ibm/sec)

and 149 kg (200 HP) power extraction. The inlet performance included in the

installed engine performance by GE was that of the Douglas Aircraft Company's

axisymmetric external compression inlet. A sketch of the Douglas inlet is

shown in Figure 37. Performance characteristics of this inlet, including pro-

pulsion drag, are shown on Figure 38. As with the turbojet engine, the propul-

sion drag as shown on Figure 38 was applied as an increment to airplane drag

polars rather than charging it as a penalty to the engine performance.

The external compression inlet was not used on the SSXJET III, instead a two-

dimensional inlet was sized to the scaled capture area of 0.814 m2 (8.766

square feet) of the Douglas inlet and fitted to the SSXJET Ill airplane. The

installed engine performance was not altered because of this change in inlet

since it was assumed that a two-dimensional inlet could be designed to produce

the same performance characteristics.

Engine performance as provided by GE was insufficient to perform mission, take-

off, landing, or noise studies. The data supplied is shown in block form on

Figure 39.

It was necessary, therefore, to expand on the data supplied to provide enough

data points to conduct the above studies. The supplied data was expanded by

first correcting it for altitude, Mach number, and atmospheric conditions, then

by adjusting the corrected data points for small changes in altitude and/Or

Mach number. The data as enriched by Vought Hampton for mission, takeoff,

landing, and acoustic analysis is shown in block form on Figure 40. Engineperformance data used in the SSXJET Ill mission studies is shown for standard

atmospheric conditions in Figures 41 through 43.

26

Engine performance data for takeoff, landing, and noise studies were expanded(see Figure 40) to cover an altitude range from 0 to 1219 m (4000 ft) and aMach number range from 0.0 to 0.4 based on the following assumptions:

i. The performance varies linearly with altitude at all Mach numbers.2. The slope of the altitude variation is the same for all po_er settings

at all Mach numbers and was determined by the full power data pointssupplied.

3. The performance varies linearly with Mach number at all altitudes.4. The slope of the Mach number variation is the same as that determined

for the GE21/J11-B10 engine at corresponding power settings of 50,48, and 46, and the same as the slope of power setting 46, for allpower settings below 46.

Graphs oF the takeoff performance of the GE21/J11-B10 engine are provided onFigures 44 through 51 for standard +lO°C atmospheric conditions.

The GE21/J11-B10 engine when installed in the SSXJET III airplane results in anaircraft design gross weight of 36076 kg (79545 Ibm). Tabulated below for theSSXJET III at various conditions are the design sea level static thrust andairplane thrust-to-weight ratios.

Conditions Thrust Thrust/Weight

Uninstalled standard dayInstalled standard dayiInstalled standard +8oc dayInstalled standard +10oc day

96526 N (21700 Ibf)90228 N (20284 Ibf)86416 N (19427 Ibf)86060 N (19347 Ibf)

.5456

.5099

.4885

.4864

Summary

From the results of these studies it can be concluded that the engine designs

provided are adequate to meet the requirements of a feasible supersonic execu-tive jet airplane. During the course of this study, no attempt was made tooptimize the engine or inlet or airplane or in combination; therefore, thefollowing are recommended:

o Conduct a detail design study of a variable cycle turbofan and/or avariable turbojet to achieve the best compromise between takeoff noisesupersonic cruise, and supersonic cruise performance.

o Conduct detail design engine inlet integration studies to provide therequired engine airflow with minimum losses.

o Conduct airplane/engine optimization studies to determine the bestengine cycle/airplane confiQuration to meet the requirements of themission and takeoff noise.

27

LOW-SPEED AERODYNAMIC CHARACTERISTICS

The untrimmed low-speed aerodynamic characteristics for the SSXJET and SSXJET Icnnfigurations have been estimated for trailing-edge flap deflections of 0°,50 , 20 o , and 30o . The lack of experimental data for the SSXJET series wing inthe high-lift configuration has necessitated the use of analytical, experimentaland empirical methods and data. This approach includes the use of a vortexlattice method for definition of the clean-wing lift curve and empirical esti-mates of landing gear and nacelle-fuselage interference drag increments.Additional data required for the analysis has been obtained from wind tunneltests of a supersonic cruise concept. These data represent the culmination ofan extensive research program employing the SCAT 15-F configuration as developedand tested at the NASA Langley Research Center. The purpose of this section isto discuss the techniques employed in the analysis and to present the resultsobtained. Although no detailed analyses have been performed for the SSXJET IIand SSXJET III, low-speed characteristics for these configurations may beestimated through suitable extension of the SSXJET and SSXJET I results as noted.

High-Lift System Definition

The high-lift system applied to the SSXJET wing is illustrated in Figure 52a.The leading-edge devices L1 and L2 and the trailing-edge flaps TI, T2, and T3are all independent plain flaps with the areas indicated in the figure. Notethat full-span trailing-edge flaps are defined for the SSXJET as the nacellesare located on the aft fuselage. The trailing-edge flaps are used in the normalsense to provide additional lift during takeoff and approach, and flap T3 mayalso be used for roll control as required. The leading-edge device LI is usedprimarily for suppression of the vortex shed from the wing apex. This vortexsuppression results in significant improvements in the longitudinal stabilitycharacteristics. Similarly, flap L2 delays flow separation on the outboardwing panel and further retards the onset of pitchup due to tip stall. Deflec-tion of these leading-edge devices does not eliminate pitchup, but it is delayedto angles of attack well beyond the normal operating envelope.

Insofar as the SSXJET I high-lift system is concerned, the primary effect ofrelocating the nacelles to the underwing position is the reduction in area forflaps TI and T2 as shown in Figure 52b, Leading-edge devices L I and L2, as wellas trailing-edge flap T3 are unchanged by the nacelle relocation. Total trailing-edge flap area for the SSXJET I is approximately 25 percent less than that forthe SSXJET. The overall aerodynamic effect of this flap area reduction is, ofcourse, of primary interest in the following analysis.

Data Base

The SSXJET is a derivative of the Douglas Mach 2.2 AST concept (reference 4).No low-speed wind tunnel data for the Douglas concept is available, however,and thus the choice of suitable data on which to base the analysis becomes the

28

fundamental concern. A survey of recent AST low-speed wind tunnel tests hasindicated that the data of reference 10 would provide the _st reasonable database. A photograph of the model used in these tests in the Langley full-scaletunnel is shown in Figure 53. The configuration has an arrow-wing planform withan inboard• sweep of 74o, a midspan of 70,5 o, and an outboard sweep of _0o.., Bothleading-edge and trailing-edge flaps are included for lift attgmentatien andpitchup control. The wing twist and camber have been optimized for IQ cruiseat Mach 2.7, and the fuselage area ruling results in minimum configuration wavedrag at cruise The model configuration included either two upper surface mountednacelles or four underwing nacelles, a drooped nose for improved pilot visibilityduring takeoff and landing, vertical wing fins for improved directione] stabilityduring cruise, and an empennage. Of particular interest are the full sDa_trailing-edge flaps tested in conjuction with upper surface mounted enginesand the increased flap effectiveness obtained by blowing at the flap knee. Theprimary data base for the SSXJET analysis is thus defined as follows:

Full Scale Tunnel Test 369 (NASA TM X-72792)

Run tl = t2 = t3 C_

320 0° O.

331 20 o .025

352 300 .025

where C_ is the flap blowing coefficient based on the wing gross area.

All of these runs have:

o two upper surface mounted engines (power off)o full span trailing-edge flapso tail offo L1,2 = 30 o (apex leading-edge flaps)o L6 : 45 o (tip panel leading-edge flap)o t 4 = 50 (tip panel trailing-edge flap)

Other data have been used where required for specific lift or drag incrementsnot directly obtainable from the above. Such departures are noted as requiredin the following text. Also note that "T" refers to SSXJET trailing-edge flapswhile "t" corresponds to test model flaps.

Lift Development-SSXJET

The initial step in the development of the SSXJET lift curves involveddefining a clean configuration lift curve to which the various lift incrememtsdue to flaps could be added. This baseline lift curve for the SSXJET has beendeveloped using the vortex lattice method described in reference II. Data inputto this program included definition of the twisted and cambered SSXJETwing,the fuselage planform, and the horizontal tail at zero incidence. The horizontaltail is assumed fixed at zero incidence throughout this analysis.

29

Development of the lift increments due to trailing-edge flap deflection requiredthat the lift data from runs 331 and 352 described above be corrected to elimi-

nate the direct lift developed by the flap knee blowing. The net lift on themodel is

= - C sin (_WRP + 6FLAP)CLNET CLTEsT _REF (1)

where C_REF is blowing coefficient based on the wing reference area.

The lift increments due to model flap deflections of 200 and 300 (relative to

0o) are then directly obtained for a series of angles of attack.

The flap lift increment correction technique described in reference 12 has been

used to estimate the effect on lift of flap geometry differences between flaps

t1, t2, and t3 for the Test 369 model and flaps T1 and T2 for the SSXJET.These geometry differences have been found to have a relatively small effect

on the flap lift.

Flap T3 on the SSXJEI corresponds to flap t4 on the Test 369 model. Insofar as

no data are available from Test 369 for t4 deflections, data from reference 13was selected for determination of the desired increment (T3 = 0o to 50). The

average value of this lift increment has been found to be .0040. This result

has been corrected for flap geometry differences between the SCAT 15-F model

reported in reference 13 and the SSXJET using the method of reference 12.

The effect of the SSXJET leading-edge device LI is to spoil the vortex liftgenerated by the wing apex. Insofar as the vortex lattice method used to

develop the baseline SSXJET lift does not account for this additional vortex

lift, the baseline lift may be assumed to reflect a deflection of 300 for LI.

The lift contribution related to a 450 deflection of L2 becomes significantonly for angles of attack beyond twenty degrees and thus has been assumed

negligible herein.

Overall lift curves for the SSXJET have been developed using the baseline lift

curve with its implicity LI and L2 effects and the flap lift increments des-cribed above. The results of this analysis are shown in Figure 54. Note that

the following configuration components are fixed:

LI = 30o L2 = 45o T3 = 5o

it = 0° (fixed)

Also note that only out of ground effect lift curves are shown insofar as

current takeoff analysis routines internally computes ground effects using the

method of reference 12. SSXJET characteristics in ground effect are presentedlater in this section.

30

SSXJET I

Overall lift curves have been derived for the SSXJET I using tile same assumptions,data base, and vortex lattice method which have been described above for the

SSXJET. The reduction in area for flaps TI and T2 has required application ofthe flap correctiom factor described in reference 4. This correction has indi-cated that approximately 60 percent of the flap lift increments derived for theSSXJET from the Test 369 data could be applied to the SSXJET I.

Nacelle-wing interference results in increased overall lift for the SSXJET I.Data from reference 13 have been used to estimate this interference lift incre-ment as follows: Lift curves with nacelles on and off have bee_ examined todetermine an average lift increment for the SCAT 15-F with four un_erwing nacelles.This increment has subsequently been corrected to account for both the reductionto two nacelles and the ratios of nacelle inlet area to wing reference area forboth the reference 13 model and the SSXJET I. This technique has resulted in

a SSXJET I wing-nacelle interference lift coefficient of .0110.

Analysis of the leading-edge devices L 1 and L2 and the trailing-edge flap T3parallels exactly the method above for the SSXJET.

The results of the SSXJET I lift analysis are shown in Figure 55. Note againthat L I = 30 o, L2 = 45 °, T3 = 5o , and that the horizontal tail is fixed at zeroincidence.

Drag Development - SSXJET

The total drag for the SSXJET is assumed to consist of pressure drag, skinfriction, interference drag,-air conditioning and propulsion bleed drag, dr_gdue to lift, and, where appropriate, landing gear drag. The development ofthese various drag items has necessarily involved the use of analytical, experi-mental, and empirical data as described below.

Skin friction and pressure drag have been estimated using the DATCOM methoddescribed in Section 4 of reference 14. Turbulent flow over a smooth, flat

plate has been assumed. The pressure drag contribution is considerably lessthan the friction drag being proportional to (t/c) 4 where t/c is the wing thick-ness ratio.

Wing-body interference is accounted for directly in the computation of the skinfriction and pressure drag items when using the method of reference 14.Nacelle-fuselage-empennage interference has been estimated using unpublisheddata for the Boeing 727 aircraft which has a nacelle enpennage arrangement verysimilar to that of the SSXJET. No other interference drag items have beenconsidered.

Air conditioning and propulsion bleed drag increments have been derived usingdata from reference 12. These previous data have been corrected to account forthe SSXJET engine airflow, a reduction from four engines to two, and the wingreference area change to the SSXJET value of 89.65 m2 (965 ft2)o

31

Semi-empirical estimates of the landing gear drag have been madeusing datafrom reference 15. The nose gear consists of two wheels mounted on a singlestrut while the main gear involves two wheels mountedon each of two struts.

The overall zero-lift drag build up for the SSXJETis summarizedin Table V.These data are applicable to the takeoff, approach, and landing segmentsofthe flight envelope.

The drag contributions of the SSXJETleading-edge devices L1 and L2 have beenincorporated as a shift in the minimumdrag point. Data from runs 230 and 233of the Test 369 data base (reference 10) directly provide the drag incrementdeveloped by the apex flaps whendeployed to 30o deflection. This incrementhas been corrected for the ratio of leading-edge flap area to wing gross area.This estimated drag shift due to LI at 30o is .00194. The increment for L2deflected 45o has been similarly derived using data from reference 16. The resultfor LI is .00435. Analysis of data from reference 13 for flap t 4 indicatedthat a 5o deflection of SSXJETflap T3 would have a negligible effect on thedrag.

Determination of drag due to lift for the SSXJETutilizes the sameprimarydata base noted above for the lift analysis. As in the lift analysis, thedirect thrust effects due to flap blowing must first be removed:

= + C cos (_WRP+ _FLAP)CDNET CDTEsT _REF(2)

These corrected data have been used to calculate polar shape parameter (PSP)values for the Test 369 model as follows:

CDo% - (CD - CDMIN)

CDo% - CDIo0%

PSP =

(3)

The 0 percent suction lines for the Test 369 model are determined from theequation:

CDo% = (CL - ACL) tan (_WRP " SO)(4)

while the 100 percent values are given by:

CDIo0 % = (CL - ACL)2/ _AR(5)

where AR is the wing aspect ratio based on the wing reference area. In these

equations CL is the total thrust corrected lift coefficient, AC, is the liftcoefficient at the minimum drag point, _WRpis the angle of attack of the wing

reference plane, and s 0 is the angle of attack for zero lift. (C D - CDMIN ) must

be determined from the thrust corrected drag data where CD is the total drag

32

coefficie_,t and CDMIrI is the minimumdrag. The drag polars to be developedfor the SSXJETwill thus reflect the "suction" levels developed by the Test369 model and will be con._.tructed about the minimum drag point.

The results of these calculations are presented in Figure 56. Polar ShaDeParameter values for the case tl = t2 : t3 = 5o were estimated by linearly

interpolating the PSP values at 0° and 20o .

The PSP values determined above have been assumed to be directly applicable tothe SSXJET. No corrections for leading-edge Reynolds number have been appliedto these PSP values. Although the direct lift and thrust effects due to flapblowing have been extracted from the Test 369 force coefficients, it has bee1_assumed that the additional induced circulation due to blowing results in

sufficiently improved leading-edge performance to justify omission of theleading-edge Reynolds number correction. Thus it becomes possible to compute

the desired (CD - CDMIN) values"

CD : - PSP - ) (6)- CDMIN CDo% (CDo% CDIo0},

where CDo% and CDIo0 % must be calculated for the SSXJET using the lift curves

from Figure 53 and the aspect ratio AR = 1.84. The ACL values required forthese calculations are not directly available for the SSXJET, and thus thevalues determined from Test 369 have been assumed. This choice of AC. valuesalso implicitly accounts for the polar vertical shift due to deflection of

L I, L2, and T 3 insofar as the corresponding model components for Test 369 weresimilary deployed.

The required lift coefficients for the polar definition have been obtained byadding to the CL - ACL values from Figure 54 the increments in ACL associatedwith the various flap deflections. The &CL value for T I = T2 = 5o has beenestimated by linearly interpolating the 0o and 20 o flap data.

Total drag coefficients have been derived by adding the zero-lift drag incre-

ments previously discussed to the CD - CDMIN values determined from equation

(6). Also included are the shifts in minimum drag due to LI, L2, and T 3 aswell as the increments due to flap deflections from Test 369. The estimatedlanding gear drag has also been included for the TI = T2 = 20 o and 30 o cases.

These drag polars for the SSXJET are presented in Figure 57. Note again thatno ground effects are included and that the horizontal tail is fixed at zeroincidence. The lift-to-drag ratio values obtained from these polars are shown

in Figure 58.

33

SSXJETI

Skin friction, pressure drag, air-conditioning and propulsion bleed drag, andlanding gear drag values for the SSXJETI have been estimated as describedabove for the SSXJET,and, with the exception of the wing-nacelle interference,the values of these zero-lift drag items correspond to those of the SSXJET.As previously noted in the lift development above, data from reference 13 havebeen used to estimate the nacelle-wing interference for the SSXJETI. Nacelleon and off data from reference 13 have been used to derive the nacelle drag forthe SCAT15-F with four underwing nacelles. An estimate of the nacelle frictiondrag has been subtracted from this increment to obtain the interference drag.This result has subsequently been corrected to account for the reduction fromfour to two nacelles and the ratios of nacelle inlet area to wing referencearea. The resultant wing-nacelle interference drag coefficient for the SSXJETI is .00275. This relatively high interference increment results because theSSXJETI nacelles are large comparedto the wing. Table VI summarizesthecomplete zero-lift drag build-up for the SSXJETI.

Drag values for leading-edge devices LI and L2 as well as trailing-edge flapT3 corresponds to the values previously estimated for the SSXJET. Drag-due-to-lift for the SSXJETI has been determined using the polar shape parameter (PSP)technique and data base discussed above. To account for the effects of thenacelle flap cutout, the following technique has been applied: Additional PSPvalues have been determined from runs 233, 239, and 251 of Test 369 (reference10). These runs provide data for the test model configured with four underwingnacelles and two trailing-edge flap cutouts. All other configuration componentscorrespond to those described above for the SSXJET. Thesedata have beencorrected for flap blowing direct lift and thrust effects and the PSPvaluesdetermined for trailing-edge flap angles of 0o, 20o, and 30°, The SSXJETIPSPvalues presented in Figure 59 have been computedby averaging these twosets of Test 369 PSPvalues. This approach is intended to provide a reasonableestimate of the PSPvalues for the SSXJETI with its single flap cutout whilemaintaining consistency in the data base.

The SSXJETI drag polars have been constructed from the various drag itemsdiscussed above. Flap lift and drag increments used in the polar developmentare the geometry corrected values. The wing-nacelle interference lift and draghave been applied as shifts in the minimumdrag point. These SSXJETI dragpolars are presented in Figure 60 where the landing gear drag has been includedfor the 20-degree and 30 degree flap cases. The L/D curves derived from thesepolars are shownin Figure 61.

The analyses outlined above suffer from a lack of low-speed wind tunnel datafor the Douglas wing on which the SSXJETconcepts are based. It is felt,however, that the results presented represent a reasonable estimate of thelow-speed aerodynamic performance of the aircraft. The analyses of the SSXJETI, in particular, represents a consistent extension of the SSXJETresults andshould provide a clear indication of the differences in the low speed charac-teristics of the two concepts.

34

SSXJETCharacteristics In Ground Effect

As previously noted, current takeoff an_ly_is routines internally computeground effects on lift and drag and thus the out of ground coefficients presentedabove have been used directly in the takeoff analysis. The untrimmed in groundeffect coefficients developed in the takeoff a.nalysis are summarized belowalong with the basic approach employed. The development of these ground effectequations has been fully documented in reference 12.

The ratio of in ground lift coefficient to the out of gY'o,ind va.)ue has beer_shown in reference ]2 to be:

CL

°i =--3-- I+. C,L

l

_.+ 3_(_) ,,__-_:7 +_:_., (_ + -- h "z-TC:+___LiZ=i:/:'-T3TTAT- T_d-+-7_]-- L __,--,'-.--A]" ._] (7)

where ho is the wing height above ground, b is the wing span, and A is thewir, g aspect ratio. At points away from the ground the drae may be expreseedas"

CD = CDI_IN + (C L - CLM)(U - _o) (a)

, CLM _' _f"where C:D is the minimum drag point is the corresponding lift ._.e_l-MIN

cient, m is the angle of attack and _ is the angle of attack for zero lift" 0 " " "

The drag in ground effect may be calculated from the expression:

CDg = CDt,IIN + aLa _ (C D - CDMIN) + (G - eo)C L(c L - l) (9)

where _;i is given by"

l

1

l + 32 (b + 4 32 (u)

A drag factor is then computed using equation (8) and (9) as

D

and aoplied to the out of ground drag coefficient values.

(10)

3.5

These equations have been included directly in the takeoff analysis routinesto automatically comDuteground effects throughout the takeoff profile. Typi-cal results for the SSXJETconfiguration are presented in Figure 62. Alsoindicated in the figures are points corresponding to operating conditionsduring the ground run, at lift-off, and at the 10.7 m (35 ft.) obstacle point.The SSXJETwing span is approximately 12.8 m (42 ft.) and for ho/b values ofunity or greater, the ground effect is insignificant.

Extension to SSXJETII and SS×JETIII Concepts

Although no detailed analyses of the low-speed aerodynamics for the SSXJETIIand SSXJETIII concepts have been made, reasonable lift and drag estimates maybe madethrough suitable extensions of the analyses presented above.

The SSXJETII concepts are similar to the SSXJETdiscussed above except for thewing-body blending and forebody shaping. Thesedifferences are assumedto havea small effect on the low-speed aerodynamics and thus the SSXJETdata are appli-cable to the SSXJETII aircraft.

Drag estimates for the SSXJET111 have been madeby correcting for the differencesin zero-lift drag between the SSXJETI and SSXJETIII. Both skin friction andpropulsion bleed drag corrections have beenmaderesulting in a net drag increaseof six counts (.0006) for the SSXJETIII relative to the SSXJETI. Differencesin lift and drag-due-to-lift have been assumednegligible.

Summary

Low-speedaerodynamic characteristics for the SSXJETand SSXJETI concepts havebeen developed using a combination of analytical, experimental, and empiricalmethods and data. These studies represent a consistent approach to the develop-ment of the low-speed aerodynamic characteristics of the configurations andshould clearly point to the differences between the two concepts in the high-lift configuration. Moredetailed analyses will require low-speed wind tunneldata for the Douglas ASTwing on which the SSXJETseries aircraft are based.

SUBSONIC/TRANSONICAERODYNAMICS

The subsonic/transonic aerodynamic characteristics of the SSXJETconfigurationhave been estimated through suitable corrections to available wind tunnel datafor the Douglas Mach2.2 ASTconcept (reference 4) on which the SSXJETis based.The study covers the Machnumberrange from 0.5 to 0.95 at lift coefficientsfrom -0.04 to 0.28. The purpose of this section is to discuss the methodsemployed in the development of the SSXJETdrag polars and to note the applica-tion of the results to subsequent configurations in the SSXJETseries.

36

Data Base

The wind tunnel data used for this analysis consisted of preliminary datapackageswhich have subsequently been published as reference 4. These originaldata were fairly limited in both data range and configuration build-up, thusnecessitating a considerable amountof componentdrag prediction. Usable dataconsisted of polars for the wing-body at Machnumbersof 0.5, 0.7, 0.9, and0.95 for a lift coefficient range of O. to 0.28.

Polar Development

Extrapolation of the test polars was performed by fairing a smooth curve toreach the desired range of lift coefficients from -0.4 to 0.28 which providedthe induced drag variation with lift for the wing-body. An additional induceddrag increment equal to ten percent of CL_2/ ARwas added to account for lifteffects of the nacelles and horizontal tail. This increment is based onunpublished data for the Boeing 727 aircraft and has been corrected to accountfor differences between the configurations and represents a conservative estimateof the nacelle and horizontal tail effects.

Skin friction corrections and interference draQ values were computedusinQ themethod of reference 18. This method assumes fully turbulent flow over a smooth,insulated flat plate and makes suitable corrections for supervelocity effects,pressure drag, interference effects, excrescences, and surface roughness.

The propulsion bleed and air conditioning drag increments for the SSXJET wereobtained from reference 12 and corrected to account for the engine size changeand reduction from four to two engines. These drag increments are illustratedin Figure 63.

Zero trim drag has been assumed throughout the analysis.

Results

The resulting subsonic/transonic drag polars for the SSXJET are summarized inFigure 64. Note that for clarity the curves have been shifted to the right asindicated by the scale in the lower left corner of the figure.

Extension to Subsequent Concepts

The drag values derived above for the SSXJET have been assumed applicable toall subsequent concepts except for the SSXJET III. In this latter case, cor-rections have been made to account for the differences in skin friction dragassociated with the increased wing size and inlet/nacelle arrangement. The

37

propulsion bleed drag has also been modified to reflect the GE2]/J11-B10enginesincluded in this concept. For the Machnumberrange under consideration, theseengines require a constant propulsion bleed increment of .00053 instead of thev_!ues p_eviously presented in Figure 63. No air conditioning drag penalty isrequired for this engine. Resulting drag poiars for the SSXJET If! are pre-sented in Figure 65.

Summary

SSXjET drag poiars for the Mach number range 0.5 to 0.95 have been derivedtkrough sLiitable correction of available wind tunnel data. No detailedanalyses for subsequent configurations have been conducted except for theSSXJET !If. The SSXJET polars derived herein are assumed applicable to theSSXJET I and SSXJET Ii concepts while the necessary corrections for theSSXJET ill have been developed.

SUPERSONIC AERODYNAMICS

Lack ef appropriate experimental data for the supersonic Mach number rangerequired by the SSXJET series aircraft has resulted in application of severe!analytical techniques for the calculation of configuration lift and drag char-acteristics for Mach numbers 1.1 to 2.2. These methods (references 18--21)have been used extensively for supersonic aerodynamic studies, and numerouscorrelations with wind tunnel data have established the validity of the aDprcach(see references 19 and 22-24). Tile purpose of this section is to describe _hemethods of analysis employed, to present further correlations of the theoreticalestimates with wind tunnel data for the Douglas Mach 2.2 AST corcept on whichthe SSXJET series is based, to discuss the analyses of the various SEXJET con-cepts, and to assess the accuracy of the various analytical methods employed.

Supersonic oneration of the SSXJET series aircraft over populated areas hingeson the level of sonic boom overpressure generated by the configurations, Therelatively lower cruise Mach number and lower weiehts as well as the high cruisealtitude might result in overpressure levels sufficiently low to allow trans-continental flight. An analysis has been conducted to assess the expectedoverpressure levels for the SSXJET series aircraft, Only "first cut" estimatesof the overpressure levels have been made and no detailed analyses have beenperformed.

Analysis Methods

The supersonic aerodynamic characteristics of the various SSXJET conce_ts ,_,ave'Ibeen estimated using a series of computer programs available at the NASAi_angley

Research Center. A brief description of these programs is given below, anddetailed information may be found in the references.

3_

Skin friction... Configuratiori skin friction coefficients have been determineJusing the Sommer and Short T' method described in referenc_ ],!_.

Tile aircraft is separated into its various components as shown in Fiqtlre 6,_ andthe individual wetted areas and reference lengths determined. ((;mp.-_en;s :suchas the wing which may exhibit large variations in reference length ,;_re further

• t h'_' COII.O__dsubdivided into strips as shown Skin friction coefficients are ....=,,assuming a fully turbulent boundary layer over a smooth_ adiabatic-,,,all flatplate with transition at the leading edge of each configuration compone_'_t. Theprogram accepts as input component wetted areas and reference l_-_,n*_,_,:,,_,.,,.,,ant;' com.-putes the skin friction drag coefficients for a given Mach numb,_,_-altituc:_, com-bination or for specified wind tunnel conditions.

Zero-lift wave draq.- The far-field wave drag program uses the supersL_lic arearule concept to compute the zero-lift wave drag of e.n arbitrary con,_guration.As described in reference 19, the program establishes a series of equivalentbodies of revolution by passing planes inclined at the Mach angle M through theconfiguration for several different aircraft roll angles 8. This procedure fordeveloping the equivalent body area distribution is illustrated in Figure 67.The wave drag of each equivalent body is determined from the yon Karman slenderbody theory which relates the wave drag to the freestream conditions an_ theequivalent smooth body area distribution. The discrete equivalent body wavedrag values are then integrated to obtain the configuration overall wave drag_

Also included in the wave drag program are a series of routines for optimizationof the configuration wave drag through suitable area ruling of the fuselage.Arbitrary restraint points may be specified at fuselage stations where minimumarea conditions already exist. The program locates these points on the averageequivalent body area distribution for the complete configuration and then solvesfor the fuselage shape giving minimum configuration wave drag while simultaneouslysatisfying the specified restraint conditions.

Wing analysis.- The winq lifting characteristics and drag due to lift have beencomputed using the method described in reference 22. Based on linearizedsupersonic wing theory, the method breaks an arbitrary planform wing into amosaic of "Mach-box" rectilinear elements which are assumed to lie approximatelyin the horizontal plane as shown in Figure 68a. This sketch is illustrative onlyin that in actual practice many more grid elements would be used. These gridelements are then employed to numerically evaluate the linear theory integralequation which relates the lifting pressure at a given field point to the wingsurface shade in the region of influence of that field point. The overall forcecoefficients for the camber surface at its input incidence are obtained by inte-grating the computed pressure distribution over the wing surface. This solutionis combined using a suDerposition technique with a flat-wing solution per unitangle of attack• The nacelle-wing interference effects are computed as describedbelow and incorporated into the wing drag polar.

39

Nacelle interference effects.- Nacelle-wing interference effects have beendetermined using the method outlines in reference 21. The program uses modifiedlinearized theory to compute the loads imposed on a warped wing surface bynacelles located either above or below the wing. A typical nacelle-wing arrange-ment is shown in Figure 68b where the shaded areas represent the interferenceregions. A fundamental restriction of the program requires that the region ofinterference from the nacelle not extend forward of the wing leading edge oroutward of the wing tip. In such cases, the disturbance field can spill overonto the upper (or lower) wing surface so that the field point pressure isinfluenced by both the camber surface and the nacelle. The interference loadscannot be determined using this method when these conditions exist. Thisdifficulty has been overcome through recent unpublsihed NASA/LRC modificationsto the basic method described above. This revised technique generates and usesan upwash field in conjunction with a wing camber surface force analysis schemeto compute the interference loads. This revised method is apDlicable to thelower supersonic Mach numbers at which the original approach could not be user1.

Validation Of Analysis Methods

The analytical approach to the SSXJET analyses presented herein was originallychosen because of a lack of supersonic experimental data for the Douglas ASTconcept on which the SSXJET series is based. Such data has recently been pub-lished, however, and thus it is possible to further validate the SSXJETanalysis methods through correlation of analytical and experimental resultsfor the Douglas transport model. Both wing-body and wing-body-nacelle caseshave been chosen and are discussed below.

Model description.- The wind tunnel model analyzed is a .015 scale representa-tion of the Douglas Mach 2.2 AST concept and is fully detailed in reference 4.A photograph of the model in the 9 x 7 tunnel at the NASA/Ames Research Centeris shown in Figure 69. The particular components involved in the analysis aredefined as follows:

B1 Fuselage for the McDonnell-Douglas D-3230-2.2-5E super-sonic cruise aircraft configuration. The model fuselageis accurately scaled forward of the full scale station2450 inches and is distorted aft of that to accommodate

the sting.

W2 - Baseline wing for the Douglas concept with the cambersurface optimized with a pitching moment constraintdesigned to give minimum drag due to lift for the trimmedconfiguration.

The wing-body (BIW2X 2 + Tbl 6"4) and wing-body-nacelle (BIW2X2N2dlcdlD + Tbl 6"4)

configurations are subjects of the present investigation.

4O

Analysis and ;'esults.- The analytical methodspreviously described have beenused to e_tim_te the model aerodynamic characteristics fo_' Math numbersof 1.6,]..8_ 2.0, and 2.2. l.ift analysis has been conducted for the twisted and cam-bered _ing alone with a constant lift displacement added equal to the nacellenormal force load on the wing as calculated using the n_ce",le interference prc.-gram. The total drag is assumed to consist of skin friction, wave drag, anddrag due to lift. Note again that nacelle effects on che _,.ring drag due to liftare included directly in the method, and that no roughness has been included.

Lifting characteristics for the Douglas model are summarized in Figure 70. Tiieagreement between theory and experiment is very good aithoug_ the lift isslightly underestimated. This disagreement is probably due in part to t.i_elifting and interference effects of the fuselage which have not beer, includedin the analysis.

Drag polar correlations are presented in Figures 71 and 72 for the win!_-.bodyan_ wing-body-nacelle configurations, respectively. The agreement between theoryan_ the tunnel data for the minimum drag point is good as is the total drag at.the lower positive lift coefficients. At the higher lift coefficients, hog,ever,the predicted drag-due-to-lift factor results in relatively significant differ-ences between the measured and predicted drag data for most Mad. n_mber's. Thisfact is reflected in the L/Dcurves presented in Figure 73 which indicate excel-lent agreement with the data for lift coefficients in the usual range of inter-

est (0. to .i0), but which disagree near the (L/D)MA X point. The variation of

(L/D)MA X with Mach number is shown in Figure 74. The predicted (L/D)MA X values

agree fairly well with the measured values although the slope of the theoreticaldata for the wing-body case is incorrect. It should be noted that these resultsare for a configuration which is slender in the classical sense and for whichthe linear theory is well suited. The SSXJET configurations are all less slenderand thus some degradation in the accuracy of the predictions is to be expected.

The methods presented above provide reasonably accurate estimates of the aero-dynamic characteristics of the major components of this slender transport modelfor the Mach number-lift coefficient values of usual interest. Good estimatesof the configuration supersonic performance may be obtained quickly and withvery reasonable expenditure of computer effort using these methods.

Configuration Analysis

Five variations of the SSXJET concept have been consistently analyzed from thesupersonic aerodynamic standpoint. In all cases the analytical techniques pre-viously described have been used for development of the lift and drag charac-teristics. Inclusion of additional drag components are noted as required.

SSXJET.- The overall lifting characteristics of the SSXJET are summarized inFigure 75. Both lift-curve slope and lift-angle of attack data are presented.The lifting efficiency of the wing as expressed by the lift-curve slope decreasesby approximately one-third as the Mach number is increased from 1.1 to 2.2.Typical operating conditions at Mach numbers of 1.2, 1.8, and 2.2 (start cruise)are also indicated in the figure.

41

Preliminary design considerations for the SSXJETdictated a fuselage length of31.89 m (I03 feet) with a series of minimumarea stations in both the cockpitand passenger compartmentas shownin Figure 76. These minimumarea Dointswere restrained in the wavedrag optimization routines and two iterations per-formed to define a fuselage giving minimumconfiguration wave drag at Mach2.2.The resulting fuselage area distributions are shownin Figure 76 along with thearbitrary initial shape. The final fuselage shape results in a configuration wavedrag coefficient which is converged to within approximately one drag count of thesucceeding iteration. Note also that the minimumarea requirements have beenfully satisfied. This optimumfuselage has been employed in the wavedraganalysis of the SSXJET. A typical Mach2.2 equivalent area distribution plotdeveloped in conjunction with the analysis is shownin Figure 77. Both thetotal area and the contribution of the various configuration componentsareshown. The computedSSXJETwavedrag characteristics are presented in Figure78 as a function of Machnumber. The cruise wave drag coefficient is .00467.Note that, although the configuration has been optimized for Mach2.20, theSSXJETgeometry as seen by the area rule and linear theory results in a relativeminimumnear Mach1.80. The increase in wave drag at the design point probablyindicates violation of the "smooth" slender body and linearized theory assumptionswhich are fundamental to the wavedrag method. Further commentsregarding un-certainties in the wave drag analysis may be found later and in the discussionof the SSXJET II concept,

Another interesting aspect of the wave drag analysis using the far-field programconcerns convergence of the solution at a given Mach number. Both the number ofaircraft roll angles (Ng) at which the discrete equivalent body wave drag valuesare computed and the number of Mach cutting planes used in the area developmentsmust be specified when the analysis is conducted. The dependence of the solutionon the value of N8 has been found to be relatively insignificant for values ofN8 = 16 or more. As shown in Figure 79, however, the number of cutting planesspecified can introduce additional uncertainties into the analysis. As thenumber of cutting planes is increased to the program limit of NX : 100, thewave drag continues to increase. A converged solution is not obtained, however,and it becomes necessary to adopt the approach indicated in the lower portionof the figure. Here the computed wave drag coefficients have been plotted versusthe inverse of the number of cutting planes employed. The curve is then extra-polated as shown to obtain the solution for NX becoming very large. This pro-cedure should be followed for all pertinent Mach numbers to obtain the theoreti-cally converged results. This convergence study was conducted subsequent to thestudies of the various SSXJET configurations which were analyzed with NX = 50.The results for the SSXJET indicate a wave drag uncertainty at Mach 2.2 ofapproximately .0004 which amounts to less than 3 percent of the total SSXJETdrag at cruise.

A configuration wetted area summary is presented in Table VII while the variationof skin friction drag with Mach number is shown in Figure 78. Note that thefriction drag coefficient increases .00035 during constant lift coefficientcruise as the altitude increases.

A roughness drag penalty has been estimated using data previously developed inreference 12. The ratio of roughness drag to skin friction drag has beenassumed to vary linearly between the values given in reference 12 at Mach

42

numberso!: I._i ar,c'.2.7. These i;',terpo]ated ratios of roughness to frlciiondrag have bee_'_ used i_ coi_junctior, with the skin frictioi_ v-a]ues discussedat, ove to estii_ate tlie SSXJET roughness drag at the required Yiach number.-:. Theresu_t.s are r,resenteJ in Figure 78.

Camber dr_q as uomp,.Ited by the wing analysis pror_'ram is also Drc.qenter:' ingig!.:re 7C,.

br'.ag increm_.nts due to propulsion bleed and air conditioning have been esti_#atedusing d_ta from reference 12. These previous data h_ve been corrected t{) accountfor the ._-,,.JET_"_..i:glne" airflow, a reduction from the four engines, to, two, and thereference area change to 89.65 m2 (9_5 ft2). Note that _n tl,is analysis theairflow ratio correction has been applied to the _.ir conditioning d_'ag incren_nt_:o obtai_l a more realistic estimate of the penalty, Figure 8[! presents theresults of this tnalyzis.

The Dougla_ win£_ (W2 of reference 4.) which has been incorporated into the SSXJETconcept has been designed to be self-tr_mming at Mash 2.2 for a ce_ter-of-grav_tylocation at 53 percent of the mean chord and at a lift coefficient of 0.10.Insofar as these conditions approximate those for the SSXJET durir_g cruise, thehorizontal tail has been assumed to be oriented tc the local flow such tha_. zero-iift is maintained on the tail with the wing at zero pitching moment. Thisassumption has been applied to all supersonic Mach numbers, and thus nc drag--due-to-lift penalty associated with the horizontal tail has been assessed. How-ever, this assumption is good over a limited Mach number range and requires morestudy. Note that skin friction, roughness, and wave drag increments due tot.he ta_1 have been included.

Selected drag polars developed through the analysis above are represented inFigure 81. Typical operating conditions at these Mash numbers are also indi-

cated in the figure. The (L/D)MA X variation with Mach number is sllown in

Figure 82 where at the cruise. Mach number, the (L/D)MA X potential decreases

from 7.05 at start cruise to 6.93 at end cruise. As noted, the actual startcruise L/D is considerably lower than the maximum achievable performance. Thisdisparity occurs because the wing size is dictated by considerations of landingand takeoff performance which results in an effectively oversized wing at cruise.

The aircraft thus cruises at a CL less than that required for (L/D)M& X. Detailed

resizing of wing and engine could provide substantial performance improvements.

SSXJET I.- The SSXJET I configuration has been derived from the SSXJET analyzedabove by moving the nacelles to an underwing location. With the exception ofthe nacelle interference load computations, the methods used to determine theSSXJET I characteristics parallel those previously described.

A sensitivity study has been conducted to determine the effect of nacelle under-wing location on the high-speed performance. As shown in Figure 83, six caseshave been considered. The first three are for constant semispan location withvarying chordwise positions while the remaining cases vary the semispan locationat constant longitudinal coordinate. In all cases, the nacelle vertical locationhas been slightly adjusted to maintain the nacelle maximum diameter tangent tothe wing trailing edge. Data defining these cases are summarized below:

43

Case Y/b/2 X/c, %

I .30 51.2 .30 60.3 .30 69.4. .4O 61.5 .50 47.6 .20 74.

Results of this analysis are presented in Figure 84. Although Cases I and 2could be eliminated from consideration due to their inferior (L/D)MAX perfor-mance, the choice of nacelle location from an aerodynamic viewpoint is notobvious. Since the Mach2.2 (L/D)MAX values are nearly equal, the aerodynamicperformance at the below cruise Machnumbersbecomesthe primary concern. In-board relocation of the nacelles results in improved (L/D)MAX performance forlower Machnumbers, but this trend does not continue throughout the Machnumberrange of interest. Since the (L/D)MAX curves cross each other at various pointsthe configuration does not appear to be overly sensitive to nacelle locationand thus other parameters such as trailing-edge flap geometry or wing structuralweight can be considered in the choice of nacelle location without serious de-gradations in the high-speed aerodynamic performance. Rangeassessmentsforthese various cases have indicated less than a 3 percent spread in the results,and thusCase I has been selected as the SSXJETI baseline nacelle location.The reader should be cautioned that the large difference betweenL/D operatingand L/DMAX was important to this result. Opportunity exists to significantlyimprove the cruise wing-engine match which would alter the optimumenqine loca-tion. Although its aerodynamic performance is somewhatinferior to other cases,this nacelle location offers a reasonable compromisefor both aerodynamic andstructural constraints. The lift and drag characteristics for this configurationare summarizedin Figure 85 for selected Machnumbers.

Typical operating points for the selected SSXJETI configuration have been in-dicated in both Figures 84 and 85. A wetted area summaryhas been included inTable VII.

SSXJET II.- The SSXJET II is a blended wing-body concept configured for eitherside-by-side or tandem cockpit seating. Exact numerical models of these twoconcepts have been prepared and used in the analyses.

Significant difficulty was encountered in the wave drag analysis of the side-by-side cockpit version of the SSXJET II. As shown in Figure 86, the predictedwave drag variation with Mach number for the exact digital geometry definitionexhibits rather erratic behavior for Mach numbers above 1.6. Modifying thefuselage geometry to an equivalent circle representation further aggravates theproblem. The difficulty here is related to the rate of area growth in the fore-body as well as the magnitude of the local body slopes. The basic assumptionsto the linear theory employed in the wave drag method are violated and theresulting drag values are probably not accurate.

44

Relaxin[i the forebody camber for the side-by-side-cockpit cc:ncept results in amore realistic" wave drag curve as sho_,,'n, a_d these values have _een inc!uded inthe polar development,

The problems discussed above were not encountered for the SSXJET II with thetandem cockpit arrangement. The more slender forebody shape included for thispilot arrangement has resulted in area distributions more conducive: to a_alysisusing the slender body theory. The wave drag variation with Mach number fortLis configuration is also included in Figure 86.

These have drag problems have been encountered primarily in the analysis of / theSSXJEI II side-by-side concept although all of the configurations analyzed showindications of Dossible theory violations at the higher Mach numbers. Theseuncertainties in the wave drag analysis are significant in some cases (i.e.,the SSXJET If), and point to the necessity of both fully understanding thelimitations of the analysis methods and the need for supersonic test data toverify the analysis methods and provide the baseline drag levels of the configu-rations.

Wetted area data for both SSXJET II concepts are included in Table VII. Themanner in which the wing-fuselage blending has been numerically modeled requiresthat the wing and fuselage wetted areas be added if a comparison of wetted areasis to be made with the other configurations.

Overall lift, drag, and lift-to-drag ratio performance for the two SSXJET IIconcepts are summarized in Figures 87 and 88. Note that the lifting character-istics of the two configurations are assumed to be the same. Typical operatingpoints are also indicated in the figures.

SSXJET III.- The SSXJET III concept has increased wing area and fuselage lengthrelative to the concepts discussed above and employs a two-dimensional inletsystem located beneath the wing (see configuration assessment for rationale).This inlet system has been modeled as an equivalent area circular nacellelocated at an "average" vertical position. No attempt has been made to accountfor the camber associated with the "S" shape of the actual inlet. It is feltthat this representation provides reasonably correct distributions of both areaand volume for the drag analysis.

Drag values have been estimated using the methods previously discussed althoughthe use of the GE21/JII-BIO engine results in a revised total additive propul-sion drag as shown in Figure 89. Also, note that an additional drag incrementto account for such miscellaneous items as inlet-fuselage interference, locallyseparated flow, the small fuselage ventral, etc. has been calculated as 5 per-cent of the skin friction drag.

Wetted areas for the SSXJET Ill are included in Table VII while the lift anddrag characteristics are summarized in Figures 90 and 91. Typical operatingpoints are as indicated.

The begin cruise (L/D)MA X of 7.0 compares favorably with similar results for

the other configurations presented above. As with the other configurations,

45

however, the actual start cruise L/D is considerably below the maximumperformance level.

"First Cut" Sonic BoomEstimation for the SSXJET

The "first cut" prediction method of reference 25 has been used to compute thesonic boomoverpressures generated by the SSXJETconcept for Machnumbers1.2through cruise. This prediction technique is illustrated in Figure 92 whichhas been taken from reference 25. Weights at altitude typical of the SSXJEThave been used along with the Machnumber-altitude schedule commonto all ofthe SSXJETseries aircraft. An "average" shape factor has been used for theSSXJETas indicated in Figure 92.

Results of this analysis are presented in Figure 93. The relatively high weightand low altitude at the lower supersonic Machnumbersresult in overpressurelevels on the order of 86.2 N/m2 (1.8 Ibs/ft2). This level steadily decreases

and is about 52.7 N/m2 (1.121bs/ft2) at start cruise. An end cruise value ofabout 38.3 N/mz (0.8 Ibs/ft ) is indicated. As pointed out i_ reference 26,sonic boomoverpressures on the order of 47.9 N/mz (1.0 Ib/ft ) probably repre-sent the maximumboomlevel which would be acceptable to the general public.It should be noted that the values computedfor the SSXJETrepresent the peakoverpressure and do not address the associated lateral decay and cutoff distanceassociated with the sonic boom. This analysis for the SSXJET indicates thatthe potential for overland flight exists for aircraft of this class, but furtheroptimization and boom minimization (through nose blunting, for example) of theconfigurations will be required.

Summary

Supersonic lift and drag characteristics for the various SSXJET configurationshave been derived using a series of analytical techniques. The validity of the

approach has been demonstrated through correlation of theory and experiment forthe ,015 scale Douglas AST configuration on which the SSXJET series is based.Analyses of the various SSXJET concepts has indicated some uncertainties in thecomputed wave drag values due to the lower fineness ratios of these configurations.In some cases it appears that the slender body theory fundamental to the wavedrag analysis has been violated and the uncertainties are large. Resolutionof these problems will require high-speed wind tunnel data for configurationsmore similar in fineness ratio to the SSXJET series than is the Douglas transport.

Sonic boom analysis for the SSXJET h_s indicated supersonic cruise overpressureson the order of 47.9 N/m 2 (I.0 Ib/ftL). Although no detailed analyses of thepressure signature shapes and magnitudes have been conducted for the variousSSXJET concepts, it is felt that the results presented for the SSXJET provide areasonable estimate of the overpressure levels to be expected. Further optimi-zation of the configuration for sonic boom minimization could result in avehicle with boom levels acceptable for overland flight.

46

MISSIONANALYSIS

The object of Mission Analysis is to evaluate the mission performance charac-teristics of a particular aircraft configuration and determine whether it attains,exceeds, or falls short of the design mission go_Is. Since the overall designobjective is to optimize an aircraft configuration, to bring together the pro-per combination of aerodynamics, airframe, and power plant configured to themir:imum size that will achieve, but not exceed the mission _oal, mission evalu-ation plays a significant role in the sizing and configuration selectio_ pro-ces_ during aircraft design.

The design mission goal for the SSXJET aircraft was defined as a range of5926 km (3200 n. mi.) at a cruise speed of Mach 2.2 with a _ayload of 726 kg

(1600 Ibm) representing eight passengers with baggage.

Methods and Criteria

A NASA developed Long-Range-Cruise Mission program (unpublished) was used inthe mission performance evaluations. Aerodynamic, propulsion, and weight datarequired as input for the mission program were developed. The input furnishedby aerodynamics consists of drag coefficients and lift coefficients for each ofa series of Mach numbers. The power plant input consists of an engine data

package which contains gross thrust, fuel flow, and ram drag data at climb andcruise power for various altitudes and Mach numbers. Design gross weight (DGW),operating weight (OW), and payload are the mass property data required. Detailedinformation regarding these data can be found in each of the respective sections.After the aerodynamic, propulsion, weight, and fuel data have been establiShedfor use as input, an unpublished NASA/LaRC mission proqram is used to perform themission analysis based on the followinq input data and criteria.

input

o Aircraft gross weight

o Aircraft operating weight

o Payload

o Aerodynamic drag polars

o Engine performance

Criteria:

Under the mission criteria established for this study, the designmission goal for the SSXJET aircraft is defined as follows: aircraft must havethe capability of flying 5926 km (3200 n. mi.) at a cruise speed of Mach 2.2

47

while carrying a payload 726 kg (1600 Ibm) consisting of eight passengers andtheir baggageand have sufficient fuel allowances to provide the following off-nominal operation capabilities.

o 482 km (260 n. mi.) to alternate airport to be flown at bestaltitude and Machnumber.

o 30 minutes in holding pattern at 4572 m (15000 ft) altitude.

o Allowance for headwinds and off-nominal operation equal to 7 percentof trip fuel.

o Missed approach - no allowance.

Theseoff-nominal operation capabilities are based on fuel reserves establishedby FAR121.648 modified for holding altitude. Current FARreserve fuel regula-tions 9overnin_ international flight include a requirement for 30 minutes at457 m (1500 ft). During SCARstudies it was found that for supersonic aircraft,which are designed to operate efficiently at higher speeds and altitudes, in-creasing the hold altitude to 4572 m (15000 ft) would result in about a onepercent range improvementand significantly improve noise. This is the basisfor the hold altitude modification which is used in this study.

The foregoing criteria establish the mission profile on which basis the missionperformance analysis is based. Mission profiles for each of the SSXJETaircraftevaluation appear in Figures 94 through 98.

During the engine and aircraft sizing studies, engine thrust to weight ratios(t/w) ranging from .35 to .55 were evaluated. With the exception of noisecriteria, the major parameters affecting power plant sizing are as follows:

o Takeoff field length

o Safety rules during takeoff which include balanced field length andmaintaining a specified minimumrate of climb with one engine inoperative.

o Climb ceiling resulting from inadequate thrust which would prevent theaircraft from reaching optimum cruise altitude.

o AdeQuateacceleration power to attain desired cruise speed, particularlythrough high drag transonic region.

o Cruise efficiency, lowest fuel consumption.

o Determination of engine performance due to above normal ambient tempera-tures, high power extraction to operate accessory systems or airbleed for specialfeatures such as surface blowing or boundary layer control.

o Safety regulations during landing and approach climb capability with oneinoperative engine.

48

The performance of SSXJET,SSXJEI"I, and SSXJETII series are based on use ofsc_led advanced dry turbojet engines operated at standard +8oc day temperaturecondition_. Aircraft and engine sizing are based on hot day operating condi-tions. SSXJET Ill performance is based on use of GE21/J11-BI0 variable cycleoperating at standard da,y conditions.

The program analyze_ each segment of the selected mission profile and providesen-route details such as required fuel, thrust: altitude, speed, and time foreach segment. These data, along with other pertinent aerodynamic, weight, andpropulsion parameters, are a)l recorJeJ on a computer printout and are avaiiabiefo_ investiqation at any discrete point e,lenq the mission. Tl]e mission profileused in this study is composed of the fol!owi_Ig segments:

]. Takeoff fuel allowance consisting of ten minutes taxi (idle powersetting) plus .one minute at full takeoff thrust with no credit for distance.

2. Climb and accelerate in accordance with the Mach-altitude climb

schedule shown in Figure 99. The program automatically determines optimumcruise altitude for maximum range unless thrust available is inadequate toreach that altitude, and a climb ceiling is established.

3. Cruise begins at either optimum altitude or climb ceiling, lheprogram determines the Brequet range factor, V/SFC L/D, for begin and endcruise, and determines an average wange factor value which is applied over theentire cruise range.

4. Descent is not calculated by the program; however, previously deter-mined values for descent distance, time, and fuel are used as input. A descentrange of 370 km (200 n. mi.) and a descent time of 20 minutes were selected foruse in this analysis. The fuel estimate was based on 20 minutes of fuel flowat idle power at the average descent altitude and speed.

Results of the mission performance evaluation are summarized in Tables VIIIthrough XII.

Off-Design Operation

Since it is not possible to operate a supersonic aircraft at cruise speeds forall missions, particularly overland, the aircraft was evaluated for subsonicrange performance. Although no mixed cruise missions were investigated,missions with all subsonic cruise segments were evaluated to determine range.A transcontinental flight, New York to San Francisco, at Mach .95 is feasible,and it was determined that the SSXJET has a subsonic range adequate to fly thisrange at maximum payload. A hypothetical mission could be te fly from SanFrancisco to New York subsonic at Mach .95, refuel and continue from New Yorkto London or Paris at supersonic cruise at Mach 2.2 (Figure 100). A supersoniccruise mission between a city-pair, (Los Angeles to Honolulu) having a range of4074 km (2200 n. mi.) which is less than the SSXJET long-range cruise capability,but so located as to make possible an all supersonic (Mach 2.2) cruise leg.

49

An all supersonic cruise mission between the city-pair of LosAngeles-Honoluluwas investigated to determine the gross takeoff weight required to fly the4074 km (2200 n. mi.) range. The takeoff gross weight for the SSXJETis deter-mined to be 30209 kg (66600 Ibm).

PerformanceSensitivity

Recent technological advances have madeavailable a unique newtitaniummaterial processing and construction technique called superplastic forming anddiffusion bonding (SPF/DB), reference 34, which yields promising airframeweight savings through increased structural efficiency. This process isdescribed in more detail in the MassProperties section of this report. AI0 percent reduction in operating weight (OW)could be achieved if SPF/DBwere used resulting in a 9 percent increase in range. The effects of changesin operating weight on range for constant gross weight and payload are dis-played in the RangeSensitivity Diagram, Figure 101.

The effects on range were also investigated and found to be 37 km (20 n. mi.)per count of drag. The change in drag versus range curve is also displayedin Figure 101. The uncertainty in calculated wave drag can result in a rangevariation of up to 741 km (400 n. mi.).

TAKEOFFPERFORMANCE

This section presents the predicted takeoff Performance characteristics forboth a SSXJETpoweredby two advancedturbojet engines and a SSXJETIIIpowered by two GE21/JII-BIO double bypass turbofan engines. The SSXJEThasa gross take-off weiqht of 35,720 kg (78,750 Ibm) and a wing reference areaof 89.65 mZ (965 ft2_. The SSXJEThas a nominal installed T/H of 0.39on a standard day corresponding to an installed SLTOthrust of 67,577N(15192 Ibf) per engine. The gross takeoff weight of the SSXJETIII is36,031 kg (79,545 Ibm) and the wing reference area is 105.0 sq.m. (1130 sq.ft.).The SSXJETIII has a nominal installed T/Wof 0.51 on a standard day whichcorresponds to a SLTOthrust of 90,210 N (20,280 Ibf) per engine.

For supersonic transport configurations, there are no existing takeoffperformance rules; however, the subsonic transport take-off requirements setforth in FAR25 (reference 32) were used as a guide. Becauseof the highlift-off velocities of supersonic arrow-wing transport configurations, thevelocity and climb out requirements set forth in FAR25 (reference 32) are notthe limiting criteria in sizing the engines for take-off. All take-off fieldlength analyses of supersonic transport configurations are conducted on astandard +8oc day rather than standard day. The installed engine thrust levelof the SSXJETis decreased from the nominal value of 67,577 N (15,192 Ibf) to64,279 N (14,540 Ibf) per engine. Similarly the installed engine thrust ofthe SSXJETIII is reduced from the nominal value of 90,210 N (20,280 Ibf) to86,416 N (19,427 Ibf) per engine.

5O

TAKEOFFANALYSISMETHODOLOGY

The takeoff method used in this study predicts the aircraft takeoff profilefrom the start of takeoff roll, through lift-off and through climb out to aspecified altitude. The methodhas been programmedand is described inreference 33. Figure 102 shows the significant parameters required to definea takeoff profile for an arrow-wing supersonic transport configuration. Thefirst segment shown in figure 102 is the total takeoff distance obtained byaccelerating the aircraft from the start of takeoff roll to a predeterminedrotation velocity, at which time the airplane is allowed to increase itsangle of attack at a specified pitch rate. The airplane continues to accel-erate and rotate until the lift-off point is reached as shown on figure 102.After lift-off, the aircraft continues to accelerate and rotate tea ma×imumallowable angle of attack and climbs to clear a 10.7 m (35 ft.) obstacle. Theall engine takeoff field length in accordance with FAR 25 (ref. 32) is thedistance from the start of takeoff roll or brake release t_ the 10.7 m (35 ft.)obstacle as shown in figure 102. The reference 33 takeoff program requiresthe all engine takeoff field length to be input and iterates on the rotationvelocity to obtain the desired result. Also, the program provides a minimumall engine takeoff field length for a particular aircraft configuration. Fromthe obstacle, the aircraft follows an adjusted climb gradient to reach a climbout point defined by an altitude and a downrange distance from start of takeoffroll as shown on figure 102. At the start of climb out the aircraft angle ofattack is gradually reduced until the desired climb gradient is obtained. Also,during this segment of climb out, the aircraft acceleration rate must bepositive. Because of the high lift-off velocities of supersonic arrow-wingtransport configurations, the velocity and climb out requirements are not thelimiting criteria in sizing the engines. At this defined climb out point, thewing trailing-edge flaps are partially retracted to a lower angle and the enginepower is reduced to that power setting required for level flight with one engineout. After cutback, the aircraft climb gradient and velocity are maintainedconstant for the remainder of takeoff as shown on figure 102.

To compute the takeoff profile, the aircraft low-speed aerodynamic charac-teristics and the engine low-speed performance characteristics are required.The necessary low-speed aerodynamic characteristics include the variation oflift coefficient and drag coefficient with angle of attack at each wingtrailing-edge flap setting for out-of-ground effect. For the aircraft nearthe ground the lift coefficient and drag coefficient are modified usingDeYoung's arrow-wing ground effect equations from reference 12 and are shownin the low-speed aerodynamics section. These equations require pertinentaircraft data such as the wing area, wing aspect ratio, the gear height, thegear drag, and the angle of attack while on the ground.

The engine thrust characteristics are also required to compute the takeoff.The variation of net thrust with altitude and forward velocity must be defined

for full power on a hot day (standard 80 C day) to determine the minimumengine size to meet the prescribed takeoff field length from the start oftakeoff roll to the 10.7 m (35 ft.) obstacle. In accordance with FAR 25(ref. 32), the all engine takeoff field length must be increased

51

15 percent to define the balanced field length, which is the actual minimumrunway length required for safety. Thus, for the SSXJETand the SSXJETIII,where the defined balanced field length is 1931 m (6500 ft), the all enginetakeoff length is 1723 m (5652 ft).

TAKEOFFFIELD LENGTHFORSSXJETANDSSXJETIII

The SSXJETis poweredby two advanced turbojet engines which have an installedSLTOthrust of 64279N (14540 Ibf) per engine on a standard +8°C day. Theleading-edge flaps of the SSXJETL1 and L2 were set at 0.523 radians (30o)and 0.785 radians (45o), respectively, and the outboard aileron T3 was setat 0.087 radians (5o). These flap settings were maintained constant for alltakeoffs of both the SSXJETand SSXJETIII. The wing trailing-edge flaps T1and T2 were set at 0.349 radians (20o), 0.436 radians (25o), and 0.523 radians(30o) for the SSXJETand both the all engine takeoff field length (distancefrom start of takeoff roll to clear a 10.7 m (35 ft) obstacle and the balancedfield length were computedfor the standard +8°C day using the takeoff methodpreviously described. The results are listed below:

SSXJETFIELDLENGTHS

WingTrailing-EdgeFlap Angle

0.349 radians (20° )

0.436 radians (25o)

0.523 radians (30°)

All Engine TakeoffField Length

1859 m (6100 ft)

1753 m (5750 ft)

1658 m (5440 ft)

BalancedField Length

2138 m (7015 ft)

2015 m (6612 ft)

1907 m (6256 ft)

Fromthe above table it was determined that for the SSXJETto meet thebalanced field length requirement at 1981 m (6500 ft), the wing trailing-edgeflap angles must be set greater than 0.463 radians (26.5o).

The SSXJETIII is poweredby two GE21/JII-BIO double bypass turbofan engineswhich have an installed SLTOthrust of 86416 N (19427 Ibf) per engine on astandard +8°C day for the GE21/JII-BIO engines, the actual all engine takeofffield length and the balanced field length cannot be ascertained for the SSXJETIII. However, the installed T/W ratio of the SSXJETIII is 32 percent greaterthan the installed T/Wof the SSXJETand the wing loading is 14 percent lessthan the wing loading of the SSXJET. Therefore, the all engine takeoff fieldlength and the balanced field length of the SSXJETI!I will be considerablyless than the SSXJET.

52

Summary

The advanced turbojet poweredSSXJETdoes meet the balanced field lengthrequirement of 1931 m (6500 ft.) with the wing trailing-edge flaps set atangles equal to or greater than 0.463 radians (26.5°).

The GE21/JII-B!O double bypass turbofan powered SSXJETIII has considerablymore available ti_rust than the turbojet powered SSXJETand, therefore, meetsthe balanced field length requirement of 1981 m (6500 ft.).

NOISEPREDICTION

Ti!is section presents the predicted takeoff noise levels for the SSXJETandthe SSX,]ET111 as well as a discussion of the FAR36 (ref. 29) noise rules andthe noise prediction methodology.

For supersonic transport aircraft configurations, jet exhaust takeoff noiselevels are a severe problem as evidenced by the Concorde. The jet exhaustnoise level is mainly dependent on the jet exhaust gas velocity. Thus toreduce the noise levels, it is desirable to reduce the jet exhaust velocity,'which in turn reduces engine thrust. The thrust loss, however, can becompen-compensatedfor by oversizing the engines. Recently, both Pratt andWhitney (ref. 27) and General Electric (ref. 28) have conducted testsunder contracts with NASA/LewisResearch Center to evaluate noisereductions of an inverted flow turbofan engine, where the outside flow hasa higher velocity than the inside or core flow. The test results withthese coannular jets have showngains up to I0.0 dB relative to single jetswith only the high velocity flow. Thus, for the SSXJETIII which uses thedouble bypass turbofan GE21/JIIBIO engines with the inverted exhaust jetprofile, there are predicted takeoff noise levels with no coannular noisebenefits and predicted noise levels using this coannular noise benefit.

Significant performance improvement is available through application of morerecent supersonic cruise technology advancements. Advancedtitanium fabrica-tion methods (superplastic forming/diffusion bonding) and application of hightemperature composite materials could significantly reduce structural weightand initial cost. A higher compressor pressure ratio turbojet engine sizedwith a retractable mechanical suppressor should be studied. A parametricsizing study is required to understand the trades between performance, noise,and field length. Finally, advanced takeoff and landing procedures (automaticflap retraction, autothrottling, acceleration in climbout, deceleratingapproach) could significantly reduce noise beyond the values estimated herein.

There are currently no existing noise rules for supersonic transport configu-rations, however, the FAR 36 rules will be used as a guide for evaluation ofthe SSXJET and SSXJET III noise levels. The initial FAR 36 rules were set in1969 and define maximum allowable aircraft noise levels in terms of effectivenoise level (EPNL) which is a time average history of the perceived noise

53

level at the observer station. Per FAR 36 (reference 29), the take-off noiselevels are measured at two stations. The first is along the runway centerlineat 6486 m (3.5 n.mi.) from the start of take-off roll and the second is alonga sideline 463 m (0.25 n.mi.) from the centerline, where the noise level afterliftoff is greatest. Figure 103 shows the location of the prescribed measure-ment stations.

For take-off noise analyses, the engine thrust data is required at full powerand at part power on a standard +lOOC day per FAR 36 (ref. 29) rather thanstandard +8oc day required for engine sizing and take-off field length eval-uation. For the SSXJET and the SSXJET III, the all-engine take-off fieldlength was set at 1981 m (6500 ft.). If a safety problem does develop duringtake-off, the engine power levels can always be increased to full throttleand the noise rules can be disregarded. To evaluate the jet noise character-istics for a particular take-off, the jet exhaust flow properties must bedefined, including area, mass flow, velocity, and temperature. These valuesare required at several flight conditions along the take-off path. The flightconditions include aircraft altitudes, velocity, and engine thrust level.

It should be noted that the amount of take-off data supplied by the enginemanufacturers is very limited, usually three or four points, so that consider-able extrapolation is required to obtain the necessary information. Thepropulsion section presents the take-off propulsion data required.

Noise Requirements

For supersonic transport configurations, there are currently no existing noiserules; however, the subsonic aircraft noise rules of FAR 36 (ref. 29) willdefinitely serve as a guideline for supersonic transport aircraft. The initialFAR 36 rules were set in 1969 and define maximum allowable aircraft noiselevels in terms of effective perceived noise level (EPNL) as presented in thenoise prediction methodology. Figure 103 shows the location of the prescribedmeasurement stations. Per the 1969 version of FAR 36 rules (ref. 29), themaximum allowable noise limit is dependent on the gross take-off weight of theairplane as shown in figures 104 and 105. Also, per the 1969 version of FAR36, the engine power was allowed to be cutback at altitudes greater than 700feet. The amount of cutback is defined as that power required to maintainlevel flight with one engine out. For a four-engine aircraft, the amount ofcutback power can be considerably less than for a two-engine aircraft such asthe SSXJET.

In 1977, FAA modified FAR 36 by reducing the maximum allowable noise limits.In addition, thrust cutback from the initial power setting was disallowed.However, for supersonic transport configurations of arrow wing design, it hasbeen maintained that thrust cutback is reasonable and in addition, the config-uration has been allowed to change by partial retraction of the wing trailingedge flaps at the cutback point. Per the 1977 version of FAR 36, the maximumallowable noise limits are not only dependent on aircraft gross take-offweight, but also on the number of engines. The allowable noise limits for atwo-engine aircraft like the SSXJET are less than for a four-engine aircraft

54

as shownin figures 106 and 107. Thus, the two-engine aircraft is doublypenalized because it requires higher engine thrust and thus higher exhaust jetvelocities and must meet reduced noise level requirements.

Noise Prediction Methodology

The aircraft observed jet noise levels are dependent on the engine jet exhaustnozzle flow characteristics, the aircraft velocity, and the aircraf: positionrelative to the observer. The engine exhaust flow characteristics include jetarea: mass flow, velocity, and total temperature. In accordance with FAR 36(ref. 29) all take-off performance characteristics are evaluated o_I _ star:dard+lOeC day.

The take-off profile was divided into nine segments and the average engineexhaust flow characteristics, aircraft velocity, and altitude were calculatedseparately for each segment. These average properties were then employed toobtain the variation of engine source noise sound pressure level (SPL) over arange of frequency and directivity angles at a radius of 45.7 m (150 ft.) fromthe center of the exit nozzle plane, using reference 30 (Stone's Interim Pre-diction Method for Jet Noise (NASA TM X-71618)). It should be noted that Stone(ref. 30) does not include noise benefits of inverted flow coannular jetswhere the outside jet flow has a greater velocity than the center or primaryjet flow.

For an observer located on the ground, at a particular instant in time, thereis a particular directivity angle between the observer and the engine exhaustjet, and at this directivity angle, the engine exhaust jet source noise SPL'sare computed over the range of frequencies from 50 to I0,000 hz at a distanceof 45.7 m (150 ft.).

The source noise SPL's are extrapolated from the source noise distance to theobserver distance using the FAR 36 (ref. 29) correction techniques. Theseinclude effects of spherical divergence, atmospheric attenuation, extra groundattenuation, and ground reflection. Spherical divergence is the dissipationof the sound pressure over a spherical surface area, as the sound wave expandsoutward from the source. Atmospheric attenuation is the dissipation of thesound pressure by the air molecules and varies with air temperature, airhumidity, the frequency level of the noise, and the distance. The higher fre-quency noise levels are dissipated considerably more than the lower frequencynoise levels. Extra ground attenuation is thought to be due to a combinationof sound wave refraction due to win_ and temperature gradients in the atmos-phere and dispersion due to the turbulent boundary layer of the earth. Theextra ground attenuation is predicted as a function of the distance, elevationangle, frequency, and wind direction. Ground reflection is the increase inobserved noise levels near the ground due to the reflection off the ground ofindirect sound waves from the source. Ground reflection is dependent on air-craft/observer geometry, frequency, ground impedance, ground roughness, andthe wave number ratio (ratio of the speed of sound on the ground to the speedof sound at altitude). The final consideration in extrapolation of the sourcenoise to the observer is the multiengine shielding effect. This is the

55

dissipation of noise from engines which are partly shielded from the observer

by the engine noise source closest to the observer. The multi-engine shield-

ing depends on the number of engines and the sideline elevation angle.

Thus, at a particular time, the variation of SPL with frequency at the observerstation is computed. These SPL's are then added logarithimically to obtain a

perceived noise level (PNL) at the observer station. As the aircraft travels

along the flight path, both the distance between the aircraft and the observer,

and the directivity angle vary. Thus, corresponding to each time during the

take-off there is an observed perceived noise level (PNL). As the aircraft

approaches the observer location and passes by the observer location, the per-

ceived noise level increases to a maximum level (FNLMAx_ and then as the air-I

craft travels away from the observer, the PNL decreases again. The effective

perceived noise level is obtained by integrating the PNL's over the time thatthe PNL first reaches lO dB below the maximum PNL until the time the PNL last

reaches lO dB less than the maximum PNL. This integrated PNL-time level is

then divided by a time interval of lO seconds to obtain the effective perceived

noise level (EPNL) in accordance with FAR 36 (ref. 29).

Take-Off Flight Profiles for Noise Evaluation

For the noise evaluation study of the SSXJET and the SSXJET Ill, the all

engine take-off field length was set at 1931 m (6500 ft.), and the engine per-formance characteristics were defined for a standard +lOOC day. The wing

trailing edge flaps of the SSXJET were set at 0.523 radians (300) and the two

advanced turbojet engines were set at a 92 percent power setting which corres-

ponds to a SLTO installed thrust of 58,660 N (13,187 Ibf) per engine on a

standard +lO°C day.

The lift-off distance for the SSXJET is 1568 m (5144 ft.) from start of take-off roll, and the aircraft velocity at lift-off is 91.4 mps (177.6 kts). Thedistance to the 10.7 m (35 ft.) obstacle is 1981 m (6500 ft.) and the aircraftvelocity at the obstacle is 96.7 mps (187.8 kts.). For minimum take-off noiselevels of the SSXJET it was judged that the cutback point be set at an alti-tude of 529 m (1736 ft.) and the cutback distance from the start of take-offroll was 5943 m (19,500 ft.). At cutback the aircraft velocity was 117.5 mps(228.3 kts.) and the climb gradient was 0.133 radians (7.64 degrees). At cut-back the wing trailing edge flap angles were retracted from 0.523 radians(30 o ) to 0.349 radians (20 o ) and the engine thrust level was reduced from55,821N (12,549 Ibf) per engine to 47,307 N (10,635 Ibf) per engine. Aftercutback the aircraft acceleration was reduced to zero and the climb angleincreased from 0.133 radians (7.64 o ) to 0.161 radians (9.25 o ) even though theengines were throttled back. This climb angle increase can partly be attri-buted to the increased lift to drag ratio obtained by retracting the flapsfrom 0.523 radians (30 o ) to 0.349 radians (20°). The low speed aerodynamicsection of this report shows that at low velocities and thus high values ofCL , the 0.523 radians (30 o ) flaps yield higher L/D ratios than obtained with0.349 radians (20 o ) flaps. However, as the velocity increases, the CL requiredto fly decreases, and at the cutback point where the CL is 0.477, the L/D with0.349 radians (20 o ) flaps is 8.84 as opposed to an L/D of 8.19 obtained with

56

0.523 radians (30 o ) flap po_ition, Io meet the takeoff field length require-ment_ it is necessary to have high v!i_g trailing edge flap settings_ whereasto minimize the engine power levels at cutback, it is necessary to Feduce thewing trailing edge flap settings.

The minimum allowable cutback thrust is defined in the 1969 version of FAR 3_irules (ref. 29) as that power to maintain level flight with oi_e e:_gine out.For the two-engine SSXJET. this power setting is considerabIv hicher the_m fora four-engine aircraft.

The altitude at the 6,48F m (3.5 n.mi.) point is 617.6 m (2026,3 ft.) ant! theaircraft velocity is I17.5 raps (228,3 kts.). Figure I08 shows the ta,kcoffprofile of the SSXJET used for evaluating the noise characteristics°

ThF SSXJET Ill has a nominal installed thrust oF 86,060 N (19_347 IbF) perengine on a standard +lO°C day, which corresponds to an installed T/W of 0.4_6roy the 36,03i kg (79,545 ibm) gross takeoff weight. Like the SSXJET, theSSXJET IIi has all all engine takeoff field length of 1981 m (6500 ft.) on astandard +lOOC day and the wing trailing edge flaps were initially set at0.523 radians (300). The GE21/JlIBlO double bypass turbofan engi_es wereoperated at power settings ranging from 63 percent to 75 percent to reducethe takeoff noise levels. The cutback distance was set at 5944 m (i9,500 ft)from the start of takeoff roll and the cutback altitude was varied from 244 m

(800 ft.) to 363 M (2250 ft.). From the matrix of cases run: it was deter-mined that the minimum takeoff noise levels were obtained at an initial take-

off power setting of 75 percent, which corresponds to an installed enginethrust level of 64,544 N (14,510 Ibf) per engine. The lift-off distance forthe SSXJET Iil is 1594 m (5231 ft.) and the lift-off velocity is 95.7 mps196.0 kts.). At the I0.7 m (35 ft.) obstacle, the velocity of the SSXJET Illis lOl.4 mps (19619 kts.). The takeoff noise levels of the SSXJET III wereminimized at a cutback altitude of 533.4 m (1750 ft.) and the cutback dis-tance was set at 5944 m (19,500 ft.) from the start of takeoff roll. Theaircraft velocity at cutback was 126.7 mps (246.2 kts.) and the climb anglewas 0.133 radians (7.630). At cutback, the wing trailing edge flaps werepartially retracted from 0.523 radians (30o) to 0.349 radians (200 ) and theengines were throttled back from 57_208 N (12,861 Ibf) per engine to 39,738 N(8933 Ibf) per engine. After cutback, the climb angle increased to 0.142radians (8.12°). At the 6486 m (3.5 n,mi.) point, the altitude is 610.8 m(2004 ft.) and the aircraft velocity is 126.7 mps (246.2 kts.). Figure 109shows the takeoff profile of the SSXJET III which minimizes the takeoff noiselevels.

Predicted Jet Exhaust Takeoff Noise Levelsof SSXJET and SSXJET III

For the advanced turbojet powered SSXJET configuration, the jet exhaust

effective preceived noise levels (EPNL's) at the prescribed takeoff measure-

ment station were computed to be I16.3 dB at the centerline measurement

station (measurement point l) and I19.7 dB at the sideline measurement

station (measurement point 2). The noise levels for the 35,720 kg (78,750

Ibm) SSXJET are shown on figures I04 through I07 and tabulated in Table XIII.

57

Together with the reference 3 limits, it can be seen from figures 104 and 105that the SSXJET exceeds the 1969 version of reference 29 rules by 17.6 dB onthe sideline and 23.0 dB on the centerline. Figures I06 and I07 show that itexceeds the 1977 version of reference 3 rules by 27.3 dB on the centerline and

25.7 dB on the sidelines.

For the GE21/JlIBlO double bypass turbofan powered SSXJET Ill configuration,

the jet exhaust EPNL's at the prescribed take-off measurement stations were

computed to be 97.1 dB on the centerline and I05.9 dB on the sideline with no

coannular suppression effects. Based on study results by Pratt and Whitney

(ref. 27) and General Electric (ref. 28) the amount of coannular suppression

could range from 5.0 dB to lO.O dB. Thus for the SSXJET Ill, the noise levelat the centerline measurement station could be as low as 87.1 dB and the noise

level at the sideline measurement station could be as low as 95.9 dB. The jet

exhaust noise levels with and without coannular suppression for the 36,031 kg

(79,545 Ibm) SSXJET Ill are shown on figures I04 through I07 and tabulated in

Table XIII, together with the reference 29 limits. From Table XIII and figures

I04 through I07, it can be seen that with no coannular suppression effect, the

jet noise levels of the SSXJET Ill exceed both the 1969 version and the 1977version of reference 29 rules. However, with coannular suppression effects,

the noise levels are below the 1969 version of reference 29 rules and the

centerline noise level is below the 1977 version of reference 29 rules, where-

as the sideline noise level exceeds the reference 29 rules by 1.8 dB.

Predicted Take-Off Airframe Nose Levels

The airframe overall sound pressure levels (OASPL) of the SSXJET and the SSX-

JET III were computed using Fink's equation (equation l) from reference 31.

OASPL = 50 log (V/lO0) + lO log (s/h2) + I00.3 (1)

On the runway centerline at the 6486 m (3.5 n.mi.) point the values used in

equation l above are as follows:

SSXJET SSXJET Ill

C-mRs (fBs)S-mL (ft)h-m (ft)

117.5 (335.6)89.6 (965)

617.6 (2250)

126.6 (4]5.0)

I03.0 (ll30)

611.O (2004)

Using these values the airframe OASPL of the SSXJET is 67.5 dB and the air-frame OASPL of the SSXJET Ill is 69.9 dB.

Summary

The turbojet powered SSXJET exceeds the FAR 36 (ref. 29) noise rules by a

considerable margin. However, by increasing the engine size so that the

58

installed T/W is the same as the SSXJET III and employing a mechanicalsuppressor, the take-off noise levels of the SSXJET could be reduced consider-ably.

The double bypass turbofan powered SSXJET III is viable from a take-off noisestandpoint in that with coannular suppression it could meet the FAR 36 (ref.29) noise limits if reduced take-off thrust is allowed.

The current 1977 version of FAR 36 noise rules (ref. 29) should be modifiedfor supersonic transport configurations, to make the SSXJET a viable aircraft.The modifications should include not imposing stricter noise limits for two-engine aircraft than for four-engine aircraft and allowing thrust cutback ataltitudes above some minimum altitude, such as 304.8 m (I000 ft.).

The coannular jet suppression effect should be investigated to ascertain theexact jet noise level of the SSXJET III.

CONCLUDING REMARKS

A preliminary design study has been conducted to determine the impact ofadvanced supersonic technologies on the performance and characteristics of asupersonic executive aircraft. Four configurations with different enginelocations and wing/body blending were studied with an advanced non-afterburningturbojet engine. One configuration incorporated an advanced General Electricvariable cycle engine and two-dimensional inlet with internal ducting. AM 2.2 design Douglas scaled arrow-wing was used throughout this study withLearjet 35 accommodations (eight passengers). Performance results aresumarized in Table XIV. All four configurations with turbojet enginesmeet the performance goals of 5926 km (3200 n.mi.) range, 1981 meters (6500feet) take-off field length, and 77 meters per second (150 knots) approachspeed. The benefits shown for wing-body blending are within the uncertaintyto which wave drag can be calculated for this class of aircraft. Noise levelsof turbojet configurations are excessive. The variable-cycle engine configura-tion is deficient in range 555 km (300 n.mi.), but meets the most stringentnoise rules (FAR 36 1977 edition), if coannular noise relief is assumed. Allconfigurations are in the 33,566 to 36,287 kg (74,000 to 80,000 Ibm) take-offgross weight class when incorporating current titanium manufacturing technology.

While the performance results to date are encouraging, some uncertainties existmainly in the prediction of aerodynamic characteristics, which can be resolvedonly by extensive wind tunnel tests through the Mach number range. Validatedlow-speed data is vital to the establishment of low noise levels, low approachspeeds, and short field lengths. Supersonic aerodynamic data is mandatory tothe transatlantic range goal and to demonstrate confidence in the applicabilityof the methods used herein. Further detailed system integration studies tothe depth described in this status report should await the completion of plannedexperimental programs.

59

REFERENCES

l , Dutton, Donnell W., Georgia Institute of Technology "Preliminary Studiesof a Supersonic Business Jet" SAE Paper 670246, Business AircraftConference, Wichita, Kansas, April 5-7, 1967.

2. Catholic University, Department of Space Science and Applied Physics,Aviation Week and Space Technology, May l, 1967, p. 77.

3. Preceedings of SCAR Conference, NASA CP-OOI, Langley Research Center,Hampton, Virginia, November 9-12, 1976.

,

°

r

O.

,

g,

I0.

II.

i2.

Radkey, R. L., Welge, H. R., and Felix_ J. R." "Aerodynamic Characteris-tics of a Mach 2.2 Advanced Supersonic Cruise Configuration at NachNumbers from 0.5 to 2.4". McDonnel Douglas Corporation, Douglas AircraftCompany, February, 1977.

Keith, A. L., Jr.: EfFects of Variable Turbine Areas on Subsonic CruisePerformance of Turbojets Designed for Supersonic Application. NASA TND-5962, October, 1970.

Baber, H. T., Jr. and Swanson: E. E.: Advanced Supersonic TechnologyConcept AST-IO0 Characteristics Developed in a Baseline-Update Study.NASA TM X--72815, January 16, 1976.

Smeltzer, Donald B. and Sorensen, Norman E.: Tests of a Mixed CompressionAxisymmetric Inlet with Large Transonic Mass Flow at Mach Number 0.6 to2.65. NASA TN D-6971, December, 1972.

The General Electric Company: Preliminary Study Data GE21/Jll-BlO StudyA1 Turbojet Propulsion System for use in Advanced Supersonic SystemTechnology Study. Report No. R72AEG329 (Contract NAS3-16950) November,1972.

Keith, A. L., Jr.: Unpublished Computer Program developed by NASA LangleyResearch Center for computing the performance of a single rotor turbojetengine with or without afterburning.

Coe, Paul L., Jr. and McLemore, Clyde: Effects of Upper Surface Blowingand Thrust Vectoring on Low-Speed Aerodynamic Characteristics of aLarge-Scale Supersonic Transport Model, NASA TM X-72792, November, 1975.

Margason, Richard J. and Lamar, John E.: Vortex Lattice Fortran Programfor Estimating Subsonic Aerodynamic Characteristics of Complex Planforms,NASA TN D-6142, 1971.

LTV Aerospace Corp., HTC: Advanced Supersonic Technology Concept Study,Reference Characteristics, NASA CR-132374, 1973.

13.

14.

17.

18.

19.

20.

21.

22.

23.

24.

25.

Lockwood,Vernard E. and Huffman, Jarrett K.: The Aerodynamic Character-istics of a Fixed Arrow-Wing Supersonic Configuration (SCAT15-F-9898)Part I. Stability and Performance Characteristics in the Landing andTake-Off Configuration, LWP-766,1969.

USAFStability and Control DATCOM:Characteristics at Angle of Attack AirForce Flight DynamicsLaboratory, dated October, 1960, revised January,1975.

Hoerner, S. F.: Fluid DynamicDrag, 1958.

Lockwood,Vernard E.: The Aerodynamic Characteristics of a Fixed Arrow-Wing Supersonic Transport Configuration (SCAT15-F-9898) Part IA.Additional Studies of the Stability and Performance Characteristics inthe Landing and Take-Off Configuration, LWP-842,January, 1970.

Quartero, C. B.: Subsonic Parasite Drag ComputerProgram. VoughtCorporation, HamptonTechnical Center, July, 1975.

Sommer,SimonC. and Short, Barbara J.: Free Flight MeasurementsofTurbulent BoundaryLayer Skin Friction in the Presence of Severe Aerody-namic Heating at Machnumbersfrom 2.8 to 7.0. NASATN 3391, 1955.

Harris, RoyV.: An Analysis and Correlation of Aircraft WaveDrag.NASATMX-947, 1963.

Carlson, Harry W. and Miller, David S.: Numerical Methods for the Designand Analysis of Wingsat Supersonic Speeds. NASATN D-7713, 1974.

Mack, Robert J.: A Numerical Method for Evaluation and Utilization ofSupersonic Nacelle-Wing Interference. NASATN D-5057, 1968.

Middleton, Wilbur D. and Carlson, Harry W.: A Numerical Methodof Esti-mating and OptimizingSupersonic Aerodynamic Characteristics of ArbitraryPlanform Wings. J. Aircraft, Vol. 2, No. 4, July-August, 1965, pp. 261-265.

Sorrells, Russell B., III and Landrum, EmmaJean: Theoretical andExperimental Study of Twisted and CamberedDelta Wings Designed for MachNumberof 3.5. NASATN D-8247, 1976.

Mack, Robert J.: Effects of Leading-Edge SweepAngle and Design LiftCoefficient on Performanceof a Modified Arrow-Wing at a Design MachNum-ber of 2.6. NASATND-7753, 1974.

Carlson, H. W. and Maglieri, D. J.: Review of Sonic BoomGenerationTheory and Prediction Methods, J. Acoust. Soc. Ameri. 51, 675-685,(1972).

26.

27.

28.

29.

30.

31.

32.

33.

34.

Ferri, Antonio: Possibilities and Goals for the Future SST, AIAA Paper75-254, 1975.

NASACR-2628,Aerodynamic and Acoustic Tests of Duct Burning TurbofanExhaust Nozzles, Hilary Kozlowski and Allen B. Packman,Pratt andWhitney Aircraft, December,1976.

Acoustic Test of Ducting-Burning Turbofan Jet Noise Simulation, FinalTechnical Report, October, 1975, P. R. Knott et al, General Electric.

Department of Transportation, Federal Aviation Administration NoiseStandards; A_rcraft Type and Airworthiness Certification, Federal AviationRegulations, Part 36.

NASATM X-71618, Interim Prediction Method for Jet Noise, J. R. Stone,Lewis ResearchCenter.

AIAA Paper 76-526, Approximate Prediction of Airframe Noise, by Martin,R. Fink, presented at the Third AIAA Aero-Acoustics Conference, Palo Alto,California, July, 1976.

Department of Transportation, Federal Aviation Administration Airworthi-ness Standards; Transport Category Airplane, Federal Aviation Regulation,Part 25.

Unpublished ComputerProgram, TAKEOFF,to Calculate Profiles for NoiseAnalysis for 5CAR Aircraft Configuration.

Pulley, John: "Evaluation of Low-Cost Titanium Structure for AdvancedAircraft", Rockwell International Los Angeles Aircraft Division, datedOctober 22, 1976.

t '

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TABLE V

SSXJET Low-Speed Zero-Lift Drag Buildup

SRE F = 89.65 m2 (965. Ft 2)

Item

Wing-Body

(including interference}

Empennage

Nacelles (including nacelle-fuselage-empennage interference)

Propulsion bleed and air-conditioning drag

Landing gear

SSXJET total zero-lift drag:

gear up

gear down

TABLE VI

i|ing-body

(including interference)

Empennage

Nacelles (including wing-nacelle

interference)

Propulsion bleed and air-

conditioning

Landing gear

SSXJET i total zero-lift drag:

gear up

gear down

Drag Coefficient

.005652

.000764

.001073

.000533

.007678

.008022

.015700

SSXJET I Low-Speed Zero-Lift Drag Buildup

SRE F = 89.65 m 2 (965. Ft2)

Item Drag Coefficient

.005652

.000764

.003603

.000533

.007678

.010552

.018230

TABLE VZll

MISSION PERFO_NCE

MISSION: Design Supersonic Cruise Mach 2.2

NODEL NO.: SSXJEI

AIRCRAFT CHARACTERISTICS

Take-off gross weightOperating weight emptyPayload-No. Passengers

CargoTotal Weight

Wing area - reference- gross

kg (Ibm)kg (Ibm)

kg (Ibm)

_ (Ibm)(ft 2 )m2 (ft2)

Advanced Turbojet C2_ ;sea level static(std. day +8°C) installed thrust

per engine, N (Ibf)

Initial installed thrust to weight ratio

Initial wing loading - reference, N/m 2 (]bm/ft 2)

- actual N/m 2 (Ibm/ft 2)

35720

16320

8

0

726

89.65

89.65

64677.3693

3907.43907.4

(78750)(35980)

80

(1600)(965)(965)

(1454o),3693

(81.61)

(81.61)

Design Mission

OPERATING

WEIGHTS, kg (Ibm)

a FUEL

kg (Ibm)

a RANGE

km (n, mi.)aTIME

minutes

Take-off

Stare Climb

Start Cruise

End Cruise

End Descent

_axi_in

Block Fuel and Time

Trip Range

NOTES: ].

2.

35720 (78750)

317.5 (700)35403 (78050)

4343.6 (9576)31059 (68474)

10644.5 (23467)

20415 (45007)

1o2.1(zzs)20313 (44782)

]]7.9 (260)20195 (44522)

o (o)

746 (403)

5075 (2740)

370 (200)

o (o)

15525.6 (34228) ,

6191 (3343)

Taxi-in fuel taken out of reserves at destination.

C.A.B. range corresponding to block time and fuel equals trip

range minus traffic allowances as will be specified for

supersonic aircraft.

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TABLE VIII - concluded

_cdel N_.: SSXJET

_eservc Fuei Breakdown, kg (Ibm):

I. 7% Trip Fuel

_, M_ssed Approach

3. 482 km (260 n. m.) to alternate airport

4. 30 min. holding at 457 m. (15000 ft)

Total Reserve

I07_

835

3243

(2378

,.

(2930)

(1 4t)

(7 49)

Initial Cruise Conditions:

Lift Coefficient

Drag Coefficient

Lift/Drag

TSFC, kg/hr/N (Ibm/hr/Ibf)

Altitude, m(ft)

.08738

.01495

5.84

.130

15240

1.276

(50000)

TABLE IX

MISSION PERFORMANCE

MISSION: Design Supersonic Cruise Mach 2.2

MODEL NO.: SSXJET I

AIRCRAFT CHARACTERISTICS

Take-off gross weight

Operating weight empty

Payload-No. Passengers

Cargo

Total Weight

Wing area - reference

- gross

Advanced Turbojet (2)

kg (Ibm)kg (Ibm)

kg (Ibm)

(ft 2)

m2 (ft2}

;sea level static

(std. day +8°C) installed thrust

per engine, N (Ibf)

Initial installed thrust to weight ratio

Initial wing loading - reference, N/m 2 (ibm/ft 2)

- actual N/m 2 (Ibm/ft 2)

Design Mission

OPERATING

WEIGHTS, kg (Ibm)

a FUELkg (Ibm)

34927 (77000)

]5756 (34736)

8 8

0 0

726 (16oo)89.65 (965)

89.65 (965)

60878 (13686).3555 .3555

3820.6 (79.79)

3820.6 (79.79)

a RANGE

km (n. mi.)

ATIMEminutes

Take-off

Start Climb

Start Cruise

End Cruise

End Descent

Taxi-in

Block Fuel and Time

Trlp Range

NOTES: I.

2.

34927 (77000)

34609 (76300)

29920 (65962)

19759 (43562)

19661 (43345)

19543 {43085)

318 (700)

4689 (10338)

I0160 (22399)

99 (218)

188(260)

o (o)

989 (534)

4745 (2562)

370 (200)

o (o)

]5384 (33915),

6104 (3296)

Taxi-in fuel taken out of reserves at destination.

C.A.B. range corresponding to block time and fuel equals trip

range minus traffic allowances as will be specified for

supersonic aircraft.

0

35

155

20

5

180

TABLE IX - concluded

Model No.: SSXJET I

Reserve Fuel Breakdown, kg (Ibm):

I. 7% Trip Fuel

2. Missed Approach

3. 482 km (260 n. m.) to alternate airport

4. 30 min, holding at 457 m. (15000 ft)

Total Reserve

1069

" 0

1300

798

3166

(2356)

(o)

(2865)

(]759)

(698o)

Initial Cruise Conditions:

Lift Coefficient .lOl/O

Drag Coefficient .01622

rift/Drag 6.27

TSFC, kg/hr/N (Ibm/hr/Ibf) .135

Altitude, m(ft) 15240

1.320

(SO000)

TABLE X

MISSION PERFORMANCE

MISSION: Design Supersonic Cruise Mach 2.2

MDDEL NO.: SSXJET II

AIRCRAFT CHARACTERISTICS

Take-off gross weight

Operating weight empty

Payload-No. Passengers

Cargo

Total WeightWing area - reference

- gross

kg (Ibm)

kg (Ibm)

kg (Ibm)

k_ (Ibm)(ft2)

m2 (ft2)

Advanced Turboiet (2). ;sea level static-(std. day+8°C] installed thrust

per engine, N (Ibf)

Initial installed thrust to weight ratio

Initial wing loading - reference, N/m 2 (Ibm/ft 2)

- actual Nlm 2 (Ibmlft 2)

33566 (74000)15436 (34030)

8 80 0

726 (1600)89.65 (965)89.65 (965)

60878 (13686)•370 .370

3671.7 (76.68)

3671.7 (76.68)

Design Mission

OPERATING a FUEL a RANGE aTIME

WEIGHTS, kg (Ibm) kg (Ibm) km in. mi.) minutes

Xake-off

Start Climb

Start Cruise

End Cruise

End Descent

Taxi-in

Block Fuel and Time

Trip Range

NOTES: 1.

2.

33566(74000)

33248 (73300)

29215 (64408)

19626 (43268)

19528 (43052)

19410 (42792)

318 (700) 0 (0)

4033 (8892) 754 (400)

9880 (21782) 4982 (2690)

97 (213) 370 (200)

118 (260) 0 (0)

14446 (31847),

6106 (3297)

Taxi-in fuel taken out of reserves at destination.

C.A.B. range corresponding to block time and fuel equals trip

range minus traffic allowances as will be specified forsupersonic aircraft.

0

29

126

20

5

180

TABLE X - concluded

Model No.: SSXJET II

Re.,';ervc_Fuel Breakdown, kg (Ibm):

I. 7% Trip Fuel

2_ Missed Approach

3o 482 km (260 n. m.) to alternate airport

4. 30 min. holding at 457 m. (1500D ft)

Total Reserve

1003 (2211)

1277 (2815)

7B4 (1728)

3064 (6754)

Initial Cruise Conditions:

Lift Coefficient .08222

Drag Coefficient .01428

tilt/Drag 5.83

TSFC, kg/hr/N (Ibm/hr/Ibf),130

Altitude, m(ft) 15240

{l.276)

(50000)

TABLE XI

MISSION PERFORMANCE

MISSION: Design Supersonic Cruise Mach 2.2

_DEL NO.: SSXJET II TANDEM

AIRCRAFT CHARACTERISTICS

Take-off gross weight

Operating weight empty

Payload-No. Passengers

Cargo

Total Weight

Wing area - reference

- gross

Advanced Turbojet (2)

kg (Ibm)

kg (Ibm)

kg (Ibm)

_ (Ibm)(ft2 )

m 2 (ft2)

;sea level static(std. day+8°C)instal]ed thrust

per engine, N (Ibf)

Initial installed thrust to weight ratio

Initial wing loading - reference, N/m 2 (lbm/ft 2)

- actual N/m 2 (lbm/ft 2)

33566

15436

8

0

726

89.65

89.65

60878

.370

3671.7

3671.7

(74000)

(34030)

8

0

(1600)

(965)

(965)

(13686).370

(76.68)

(76.68)

Design Mission

OPERATING

WEIGHTS, kg (Ibm)

a FUEL

kg (Ibm)a RANGE

km (n. mi.)mTIME

minutes

Take-off

Start Climb

Start Cruise

End Cruise

End Descent

Taxi-in

Block Fuel and Time

Trip Range

NOTES: I.

2.

33566 (74000)

33248 (73300)

29273 (645377

19341 (42640)

19245 (42427)

19127 (42167)

318 (700)

3975 (8763)

9931 (21897)

97 (213)

118 (2607

o (o)

743 (401_I

5086 (2746)

370 (200)

o (o)

14439 (31833)!

6199 (3347)

Taxi-in fuel taken out of reserves at destination.

C.A.B. range corresponding to block time and fuel equals trip

range minus traffic allowances as will be specified for

supersonic aircraft.

0

29

128

20

5

182

TABLE XI - concluded

Model No.: SSXJET II TANDEM

Reserve Fuel Breakdown, kg (Ibm):

I. 7% Trip Fuel

2. Missed Approach

3. 482 km (260 n. m.) to alternate airport

4. 30 min. holding at 457 m. (15000 ft)

Total Reserve

Initial Cruise Conditions:

Lift Coefficient .08237

Drag Coefficient .01412

Lift/Drag 5.83

TSFC, kg/hr/N (Ibm/hr/Ibf) .130

Altitude, m(ft) 15240

(1.276)

(50ooo)

1002

1277

784

3064

(22]0)

(2816)

(1729)

(6756)

TABLE XII

MISSION PERFORMANCE

MISSION: Design Supersonic Cruise Mach 2.2

MODEL NO.: SSXJET Ill

AIRCRAFT CHARACTERISTICS

Take-off gross weightOperating weight emptyPayload-No. Passengers

CargoTotal Weight

Wing area - reference- gross

kg (Ibm)

kg (Ibm)

kg (Ibm)

(ft 2 )m2 (ft 2 )

GE21/Jll-BlO engines (2_sea level static

(std. day ) installed thrust

per engine, N (Ibf)

Initial installed thrust to weight ratio

Initial wing loading - reference, N/m 2 (lbm/ft 2)- actual N/m 2 (lbm/ft 2)

36081 (79545)

]8094 (39089)8 8

0 0

726 (1600)

104.98 (ll30)

I04.98 (If30)

90209 (20280).510 _510

3370.5 (70.39)3370.5 (70.39)

Design Mission

OPERATING a FUEL A RANGE ATIME

WEIGHTS, kg (Ibm) kg (Ibm) km (n. mi.) minutes

Take-off

Start Climb

Start Cruise

End Cruise

End Descent

Taxi-in

Block Fuel and Time

Trip Range

NOTES: I.

2.

3608l (79545)

318 (700) 0 (0)35763 (78845)

2608 (5750) 193 (I04)33155 (73095)

10547 (23252) 4812 (2598)22608 (49843)

If3 (250) 370 (200)

22495 (49593)

ll8 (260) 0 (0)22377 (49333)

13704 (30212),

5375 (2902)

Taxi-in fuel taken out of reserves at destination.

C.A.B. range corresponding to block time and fuel equals trip

range minus traffic allowances as will be specified for

supersonic aircraft.

0

9

124

2O

5

158

TABLEXII .- concluded

Model No,: SSXJET Ill

Reserve Fuel Breakdown, kg (lh_):

]. 7% Trip Fue_

2. Missed Approach

3, 482 km (260 n. m.) to alternate airport

4, 30rain. holding at 457 m. (150D0 ft)

Total Reserve

,qs_(2o97)

L w

]65_(3640)

1o)4(2367)

3676 (s_,o4)

Initial Cruise Conditions:

Llft Coefficient .07865

Drag Coefficient .0|448

Lift/Drag 5.43

TSFC, kglhr/N (lbmlhr/lbf) ,126

Altitude, m(ft) 15240

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REFERENCE

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PROPULSION SYSTEM

TYPE

NUMBER

CONDITIONS

MISSION SUMMARY

MODEt_..SSXJET

ibm,=

k9

]bin

Ibm

Ibf

INSTALLED THRUST/ENGINE. N

INSTALLED TIw

T&KE-OFF FIELD LENGTH

__ _STA_RT CRUISE ALTITUDE

Wq

Ibm

.||2

m_

tt z ,

m 2

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Basic7875057205980

163201600726

4117N18674

96589,65

96589.65

81.613907.4

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dv.Turbojet

7700034q7734736157561600726

4066418445

95589.65

79.793820.679 7q

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1981 1981 198154600 54500 563001664Z 1661Z 1/16U65200 63400 64900

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TRIP TIME

RESERVE FUEL

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rain

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31587 31573 J 2995_ ]14328 14M! J 13580 {

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= -.

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--41T2-win." ALT.AIRPORT _i

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TOTAL k_

INITIAL CRUISE CONDITIONS

CL

CD

.08738 .10170 .08222 .08237

.01495 .01622 .01428 .01412

LI D 5.84-- Ibm IhrllM I " IZ /

TSF C kg_hr_N

6.27 5.83 5.83

1,3ZU 1,Z/6 1.2/6

TABLE XIV

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ORIGINAL PAGE B

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CLIMB ACCEL ----k\_//

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0 MIN TAXI + /1

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OR CLIMB CEIL_____

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Figure 27. - NASA Turbojet engine and nacelle scaling factors.

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Z"' .0020(__

,, .0018i,

o 00160

C__

.0014

:_ .00120

c_-_ .0010r_

o .0008m,-

• 0006

0 .5 1.0 1.5 2.0

MACH NUMBER

Figure 28, - NASA Turbojet total propulsion drag increment.

16 x I0_70 x 103

STANDARD DAY +8°C

14

12

60

2O

10

5O

0

_ "%

0 00 ! !

5 I'0 I'5 20 x 103PRESSURE ALTITUDE, m

i'o 2'o ' ' ' _o '0 30 40 50 70 x 103

PRESSURE ALTITUDE, ft.

Figure 29. - Installed turbojet net engine thrust for maximum climb andcruise.

STANDARD DAY +8°C

16 x 103

7 x 103

14-

6

12-

5

"4

,..J

_.,/

"' 3

6. u-

_'I0"

Er_

o 8..J

...1

24.

2 1

I . | | ,

O" 0 0 5 I0 15 20 x 103PRESSURE ALTITUDE, m

I I ! ,; IO io ao go _o 60 _ox io_PRESSURE ALTITUDE, ft.

Figure 30. - Installed turbojet fuel flow for maximum climb and cruise.

STANDARD +8°C DAY

16 x 1037 x 10 ?,

NOTE:

14 PA = Pf',ESSURE ALTITUDE /

6 G BE,J!N CRUISE /_.-M- n

M = 2.2 PA = 16642 m (54600 ft.) / 'PA- <'0

12 _-} END CRUISE / / ' :=_ 25 M= 2.2 Z /

_ _ _ /

_ PA = 19_7._ m (65200 ft) /'_'-_./E 1 _. / /'.-M = 0.4

" :_ , / / / PA = ._048 m (10000 Ft.)

__ ,-. / / --_,= 0.8 ,,=_ _ I _// / PA= 7620.,(_000 ft.)

,,.3 ///--,'_--_--'- M: 1.o ,

....._/._" PA: l.s_.o• (sooooft..)/// _'-"--- _,- 22

1 /1/ " - "2 PA = 1828S m (60000 ft.)

/

0 0 '_-'-------_ .... _ ......... ' i ___,_0 i0 20 30 40 50 60 70 x 10 _

p. rNET EN,IIN_. THRUST, N

0 5 10 15 x 1,J,"

NET ENGINE THRUST, lbf

Figure 31, Installed turbojet fuel flow for maximum and part power cruise,

STANDARD +IO°C DAY

15 x 10 2

10

r_0

w

_- 02 ,,=,_15xl

_- lO <

i---m

'" 15 x lO 2

0

10 x 10 2

o!fI.--

lO x 10 2 M = 0

.: o2-_ _ "-,o 04M = 0.4-_ "_ PRESSURE ALTITUDE

"_M = 0 1219.2 m (4000 ft.)

M_M= 0.4

RE I UDE

-- _ _-M = 0 609.6 m (2000 ft.)L-M= 0.2

x 10 x 10 2I.i.I M : 0.4-_

M = --_ _ PRESSURE ALTITUDEM : 0 0

I I I I I I J

10 20 30 40 50 60 70 x 103

NET ENGINE THRUST, NI ! I

5 lO 15 x 10 3

NET ENGINE THRUST, Ibf

Figure 32. - Installed turbojet exhaust gas temperature for take-off andpart power cruise.

STANDARD+ lO°C DAY

I0 x 1023 x 103

_

5

0 uO

._.I0 x 102"-- 3 x IO3E

2 _C)_.J*,,5i

l m

b.-

0 _<x'r =I0 X 102"' 3 x I03_

t--

c.)0

.(.f},:C

j.-u'b

.

,

O 0

M : o.4-_,

M = 0.2-_,.__-_-M= 0

____PRESSURE ALTI TUDE-_/ 1219.2 m (4000 f_;)

M : 0.4-_

.J_._ PRESSURE ALTITUDE0

____IL L___,.._.,,_. = L I

0 I0 20 30 40 50

NET ENGINE THRUST, N! i .... I ....

0 5 I0

NET ENGINE THRUST, Ibf

L i

60 70 x 103

15 x I0 _

Figure 33. - Installed turbojet exhaust gas velocity for take-offand'part po_ver cruise.

STANDARD +IO°C DAY

l " ,l

(N

4-J (N

.3,_3-_

,_C "=C

I-" I--,,) ILl o_

I-'- I---

_ . -r .l

_ L_

.33"

.22'

L •

1 .I 0

/-M = 0.2M=O

M = 0.4_

PRESSURE ALTITUDE

1219.2 m (4000 ft)

M- 0.4-- 1 '

M = 0.2-__.

-M:o

/--M = 0.4

PRESSURE ALTITUDE

609.6 m (2000 ft)

M=O.2

Sf.;o

-/_M = 0.4M=O.2

PRESSURE ALTITUDE0

M=0.4

M 0 2_

M =.0 j

I I I ! I I

I0 20 30 40 50 60

NET ENGINE THRUST, N

L ,,, I , I

0 5 I0

NET ENGINE THRUST, Ibf

I

70 x 103

I

15 x 103

Figure 34. - Installed turbojet nozzle exit area for take-offand part power cruise,

STANDARD+ IO°C DAY

180- 80 r

1-60- l70_-

I

140 " 601fJ L_

120"_

E 180 "_ 80[e"- I

m

16o- - u_ 70_

m I

14o-

N _ 60I-- F-u_ 120 -u_"_ 180"4"" =: 80rL_J ILl

160

i40

]20

M = 0.4 --\

M- 0.2-_

M=O 7 PRESSURE ALTITUDE

1219.2 m (4000 ft)

M:O _ PRESSURE ALTITUDE

609.6 m (2000 ft)

M = 0.4 -_

M = 0.2_C

70F

i

6oIII - I , I , , I l • . I: . J

0 I0 20 30 40 50 60 70 x I0 _

NET ENGINE THRUST, N

I .... I , , ,,, | _ ._ I

0 5 10 15 x 103

NET ENGINE THRUST, Ibf

PRESSURE ALTITUDE

0

Figure 35. - Installed turbojet exhaust gas mass flow for take-off

and part power cruise.

e

I-'-

X_3V3

0

"[3OJ

%0

°_.._r_

0

If,....

r--C_J

!

(I)S-

LL.

\'\\\

!

I

I--.

\

, r..ll.

EG

o_

_.)

o..E

L)

p..--

r--'

y

d

4-

U

° t""

0

!

G_

°r-

[i..

U

0

F-

r_c

LL

C_tZ3

1.0

.9

.8

.7

A = .814 m2 (8.766 ft 2)

(a) TOTAL INLET MASS FLOW RATIO

*_ 1.00*'* (3-

C_J,_ .98

or) (3_L=J

>_" .96r_

I-- bJ'" > .94

_ .92(b) INLET PRESSURE RECOVERY

. OOO8

o _ 0006L/3 (..).._I

LLr_ LL.0 L_(_C 0_ .0004 t

0

REFERENCE AREA = 105 mz (1130 ft _)

I .. I I I I

0.5 l.O 1.5 2.0 2.5

MACH NUMBER

(c) GE21/JII-.BIO TOTAL PROPULSION DRAG INCREMENT

Figure 38. - Douglas Aircraft Co. External compression inlet performance.

"Z,._

l,--.=J

8,,"] x I C:_

F70.-

t 20

5o - ,,T 15C_

40-_--

1030

20,5

25 x I0 _DF

GE21/JII-BlO IvD)_!MUM CLIMB AND M?XIMUri CRUISE

_TAKE-OFF, LANDING AN[) NOISE DATA ENVELOPE(FULL AND PART POWER CRUIS[)

0 MAXIMUM CLIMB AND MAXIMUM CRUISE

MAXIMUM CLIMB, MAXIMUM CRUISE 0

AND PART POWERCRUISE m

--- SSXJET MACH-ALTITUDE

CLIMB SCHEDULE

!/i::il

u_/

/

0_"

.._--

/

0 0.5 1.0 I.5 2.0 2,5

<> GE21/Jil-BlO TAKE-OFF

_GE21/J!l-BlO TAKE-OFFAND PART POWER CRUISE

GE21/JII-BiO MAXIMUM CLIMB,MAXIMUM AND PART POWER

CRUISE

MACH NUMBER

Figure 30. - GE21/JII-B!O data supplied relative to minimum data required.

_GE21/JII-BIO TAKE-OFF DATA SUPPLIED

80 x

<_GE21/JII-BIO TAKE-OFF AND PART POWERCRUISEDATA SUPPLIED

[] GE21/JII-BIO MAXIMUM CLIMB AND MAXIMUM CRUISE

0325 x lO 3 DATA SUPPLIED

7O

60 ¸

4- 50

_2

_- 40

30

2O

l0

2O

E 15!i,t

F-

F--

< 10

[]JGE21/JII-BIO MAXIMUM CLIMB, MAXIMUM ANDPART POWERCRUISE DATA SUPPLIED

rm

//A

0 0.5 l.O

SSXJET MACH-ALTITUDE

CLIMB SCHEDULE Z_ N _ll[]-

Z_ _jD j

crltA

zs_ kJ _ ENVELOPEOF ENRICHEDTAKE-OFFAND PART POWER CRUISE

Z_ ENRICHED MAXIMUM CLIMBAND MAXIMUM CRUISE

& ENRICHED MAXIMUM CLIMB,MAXIMUM AND PART POWER CRUISE

I i I

1.5 2.0 2.5

MACH NUMBER

Figure 40. - GE21/JII-BIO data as enriched by Vought Hamptonrelative to data supplied,

STANDARD DAY

(sl

F--

F-

25 x 103

ZOZ

_c

_Y

IB ,..Z

Z

F-ZAJ

10 ==

5

120 x 103

100 \\ M = 0,4

\\M= 0.480

80

4O

200

MAXIMUM CLIMB

--- MAXIMUM CRUISE

".95 1.4 ,.9

| ,, . i l ._

5 lO 15

PRESSURE ALTITUDE, m

| . _| . . i . ! ! |

0 10 20 30 40 50

PRESSURE ALTITUDE, ft.

2.2

\2.2. .|

20 x IOs

|

60 x 103

Figure 41. - G_E, 21/OllBlO Installed net thrust for maximumclimb and maximum Cruise,

STANDARD DAY ORIO_VAL PAGE

L.e-

E

O-./l.,L

.-/

LI-

30 x 103

25-

e-

20-

0

LI_

15 "'

10

5

14 x 10 3

12 -_

M=0.4

lO -

\8 \

\M=0.4

6

4

20

--MAXIMUM CLIMB

----- MAXIMUM CRUISE

.6

5 l 82.0

l

_ . 1.4\ \ \ \

\ \kX\l 1 x \\\• \ ._ \\ \ \ ,,

\ \• 9\X _ \

I I I

I I0 lO

lO 15

PRESSURE ALTITUDE, m

I I I I

20 30 40 50

PRESSURE ALTITUDE, ft.

2.2

\2.2

I

20 x 103

I

60 x 103

Figure 42. - G.E. 21/JIIBIO Installed fuel flow for maximumclimb and maximum cruise•

I

Z

I--

CDr,--

x

l_e)

r--

X!

Oe----

N

II

=E:

I i |

,_- l'_ k"%!

_4/BN 'M07d 73n3

l t I

aq/uJql. 'M07-1 7317._4

I_J

t_7C,

X

0

I

O")

L_Z

Z

I

I

0

ISD

I--

Z_._

ZLO

Z

O

0_

O

r_rn

cr_

E

E

O4-

Or--4-

4-

_J

%

c--

_nr-"

r--.

C_

!

STANDARD +IO°C DAY

U

llO

lO0

90

80

70

240

220

200

180

160

140

_. 60120

o; 260,,-, 24o ,,-, 110

< 220 _: I00

200 90

< 180 _ 80X X

160 __ _ 70

_ 130

]20

280

260

240

22O

200

180

llO

lO0

90

8030

%_ %.%

• .)

_ PRESSURE ALTITUDE___ - - --

609.6 m (2000 Ft.)

_ PRESSURoE ALTITUDEi _ , i 0 i

40 50 60 70 80

NET ENGINE THRUST, NI I

I0 15

NET ENGINE THRUST, I bf

I

90 x 103

!

20 x 103

Figure 44. - Installed G.E. 21/JlIBlO primary exhaust gas

flow for take-off and part power cruise.

STANDARD + I O°C DAY

25 x 102

u 20 f

15 L uVl

4-) _n

E:2 .

2.5 X 1'02_c)

2_

',' 20

X -i-

25 x 102 <

2O

15

8x 102

_SURE ALTI TIJDE

1219.2 m (4000 ft.)

8 X 10 2

7

6

5

4

609,6 m (2000 ft,)

8 x I0 z

7

6

5

4 _--=- L ,- }

3O 4O

0

I, _ , #

50 60 70

NET ENGINE THRUST, N

! I ,

10 15

NET ENGINE THRUST, Ibf

t

80 90 x 103

!

20 x 103

Figure 45. - Installed G.E. 21/JIIBIO primary exhaust jet velocityfor take-off and part power cruise.

STANDARD + l O°C DAY

.33

.323.4

.31

3.2 .30

.29

N 3.0 - E .284-

2.33_ _ .32_ 3.4-"'

_ _ .31I-- m

N 3.2 _ .ao•-r- _× .29

_ 30 _ 28:E: r__. :- .33r_

.323.4

.31

3.2 .30

.29

3.0 .28 l30

UDE

J 1219.2 m (4000 ft.)

(b.Ix

_D _ "_

0E

609.6m (2000 ft.)

I l I I I J

40 50 60 70 80 90 x 103

NET ENGINE THRUST, N

I I a

I0 15 20 x 103

NET ENGINE THRUST, Ibf

Figure 46. - Installed G.E. 21/JIIBIO primary exhaust jet area fortake-off and part power cruise.

STANDARD + IO°C DAY

20 x 102

r,,,"0

r,,

I.i.Il--

.-J

I.--0

Cr_

l--

-e"

18

16 9

14

,.-7

I--

20 x 102_l

i..i.II---

18 -I0

I--

16 29

14 _8

I--'m7

II x 102

M=0.4"\"I0 0.2 .

PRESSURE ALTITUDE1219.2 m (4000 ft.)

x 102

18-

16

14

x 20 x 102''_ II x 102

< _10

9

8

73O

M= 0.4_

M = 0.4_

I I I i i

40 50 60 70 8O

I0

NET ENGINE THRUST, N

l

15

NET ENGINE THRUST, Ibf

90 x I03

J

20 x 103

Figure 47. - Installed G.E. 21/JIIBIO primary exhaust gas totaltemperature for take-off and part power cruise.

STANDARD+ lO°C DAY

0

I--

Z

4°f3O

2O

10 _

4° f :_

30 _'

20 __1o _-

-'r"X

Z40 ,_

30

20

10

20I15

I0

5

20E15

I0

5

M=0.4

_o _PRESSURE ALTITUDE

1219.2 m (4000 ft)

M=O.4_0.2

PRESSURE ALTITUDE

609.6 m (2000 ft.)

20 M = 0.4

_-- 0.2

15 _ -0

10 PRESSURE ALTITUDE0

I I I I I5 0 40 50 60 70 80

NET ENGINE THRUST, N

II0 15

NET ENGINE THRUST, Ibf

I90 x 10 3

I

20 x 10 3

Figure 48. - Installed G.E. 21/JIIBIO fan exhaust gas flowfor take-off and part power cruise.

STANDARD+ I0 ° DAY

20 x 102 6 x 102 M = 0.4-_

I !15

4 _RESSURE ALTITUDE_ 1219.2 m. (4000 ft)

"-,_ 10 ._.34- E

20 x 102 6 x 102 M = 0.4-_i_ i---¢

'"515 >

t-.-I,D _

10 _

x

"' 20 x

LI.

15

I0

PRESSURE ALTITUDE609.6 m (2000 ft.)

-=_

X

02 '"6 x 102 M = 0 4

'_ 0.2Lt-

5

4 _ PRESSURE ALTITUDE-0

I ,. I I l , i I

3 30 40 50 60 70 80 90 x 103

NET ENGINE THRUST, N

I0 15 20 x 103

NET ENGINE THRUST, Ibf

Figure 49. - Installed G.E. 21/JIIBIO fan exhaust jet velocityfor take-off and part power cruise.

STANDARD + lO°C DAY

cq

r,,,:c

F-

x,,i

i .036 F

.38 '0341

.36 '0321.34 •

.32 .030 L

E

.36 !

.34 .032• 32 _ .030L

xw

.o36F.38 .034[

.36.34 "032 I

.32 .030 Li30

t_= O__._----__RESSURE ALTITUDE

0_-2__1219.2 m. (4000 ft.)

M -- 0 P_{ESSURE ALTITUDE

_2 _609.6 m (2000 ft.)

l

40

M = 0 PRESSURE ALTITUDE

I l I I

50 60 70 80 90 x 103

NET ENGINE THRUST, N

l l JIO 15 20 x 103

NET ENGINE THRUST, Ibf

Figure 50. - Installed G.E. 21/JlIBlO fan exhaust jet areafor take-off and part power cruise.

STANDARD + IO°C DAY

9 x 102

r_0

mr

F--

my-LiJ

F--

..J

F--0F-

F-Ct}

8

7

6

9 x

8

7

6.

4

x

5 x 102

MO.=2 0.4\

PRESSURE ALTITUDE

1219.2 m (4000 ft)O'

3

rw

< 5xlO 2r_

l,lF-

.=J

_ 4

I.--

_ 3I---

_- 9 x 102_ 5 x 102xi,i

z

t,

4

M=0.4_

7

6

PRESSURE ALTITUDE

609.6 m (2000 ft.)

M= 0.4,

0.2

PRESSURE ALTITUDE0

3 ' , ' ' ,. t30 40 50 60 70 80

NET ENGINE THRUST, N

I

90 x 103

I I I

I0 15 20 x 10 3

NET ENGINE THRUST, Ibf

Figure 51. - Installed G.E. 21/JIIBIO fan exhaust gas totaltemperature for take-off and part power cruise.

•9L22mT- _

(9.903 ft 2) _w_//

1.5L51m2----_ J /

(a)SSXJET

8T]3m 2

667 ft 2 )

2

(25.736 ft 2)

I. 36 m2

(14.688 ft2)

L2 _ _--- T 3

.92m2_ /r7Z/ .m m2

(9.9o3f_ 1 _/_7 18.6_7f_l

1.55 mz \ / /-I '"

T1

.73 m2

(b) (7.813 ft 2)SSXJET I

Figure 52. - High-lift system geometry.

1.4

1.2

1.0

SSXJET

L I - .524 rad. (30 °)

L 2 = .785 rad, (45 ° )

T 3 = .087 tad. (5 ° )

Sre f = 89.65 m2 (965 ft2)

i t= 0 rad. (0 °) (fixed)

OUT OF GROUND EFFECT

UNTRIMMED

CL

0

-.2

-.14

|

-8

-.07 0

i I

-4 0

.07 .14 .21

_wrp - radians

I I I4 8 12

- degreeswrp

.28

!

16

.35

I

2O

Figure 54. - SSXJET lift curves.

1.4

SSXJEI I

LI = .524 rad. (30 _)

L2 = .785 tad. (45 _)

T 3 = . 0,.£7 tad, ,,,(::_'I,

= 89 65 m (965 ft )Sre f

i t = 0 tad. (0 _) ('Fixed)

OUT OF GROUND,EFFECT

UNTRIMMED

].2

l.O

CL

0

-. 07 0

1 ! , I

-8 -4 0

_wrp - radians

• .. l ,_ • _._ I i

4 8 12 16 20

- degreeswrp

Figure 55. - SSXJET ] Lift curves.

11

O

II

dsd - _13J.3NV_JVd3dVHS tlV]0d

CO

-J(.D

I

._1(..)

I'--

x

c_

0

Q;r_0r---QJ

S..QJ

E

J_

S-

O

I

_d

L

.F..-i,

II

SSXJET

LI = .524 tad. (30")

L2 = .785 rad. (45°)

T3 : .087 rad. (5°)

Sre f = 89.65 m2 (9C5 ft2)

i t = 0 rad. (0 °) (fixed)

OUT OF GROUND EFFECTUNTRIMMED

.7

CL

.6

.5

.4

.3

.2

.]

T_ = T 2

.524rad.

30 ° )

0 .02 .04 .06 .08 .I0

GEAR DOWN FOR .349 rad. (20 _)AND .524 rad. (30 ° )

A CD gear = 0.00768

.12 .14 "16 .18 .20

CD

Figure 57. SSXJET Drag polars.

0

l_l

U

f_E

0

a_,

r_

I--,

I

m

i,

iI

_o---LCJ o

X 04 03 COtr) i._ i_-- c)if)

ii ii i!

(-4 c_....I' _J I---

,_ (-,4 c_ Qo

! | I

dSd-_313WV_Vd 3dVHS _VqOd

0

r*.-,

,,_'T-,

O0

I.c)

,:::t

CD

_.1(_.)

!

__1 I--"I_1..I

>_trJor)

S.-0

4-.

S.,• _j

E

m

(L).

..Cu_

%0-

I

o4

aJ

0 0O,J0_)v

0 0 0

0

-J

0

0

0_-_

n_

L_J

I

c;

°r-

I.--"

Xc/%

C,.Jp=.=

0

t---

Z

0_rm

II

I'--

II

CO

C}

..-J

CO

0

C=

Er,-o

c3..

c_

._._i

l--"ILl

X

_0

CT_°r=-I,

r.--- ¢--.- _" _--- |"

..J

0

O_

O0N

N

o

N

O00

@

0

CON

:,=--N

0

r_0

!

i

r-x

_o

oJ

oo

__. _n --_rY

"t3

--_C_

I

CO!

U

9--9--

E30L

E.p .

U.pE

E

OS.-

_0

E--i,i

X0")LF)

I

dtD

SSXJET

TNO ENGINES

.0020

Sre f -- 89.65 m2 (965 ft. 2)

.0016

aCD .0012

.0008

.0004

P

/AIR-CONDITIONING

I I I I I0 .2 .4 .6 .8 1.0

MACH NUMBER

Figure 63, - Propulsion bleed and air-conditioning drag increments.

SSXJET

Sre f = 89.65 m2(965 ft. 2)

CL

• O8

0 0 0 0 .012

M = .5 .7 .9 .95

• 016 .020 .024 .028 ,032

CD

.036 .040

Figure 64. - Subsonic/transonic SSXJET polars.

SSXJETIII

Sref = 104.98 m2(I130 ft.2)

.5 .7 .9.95

CL

•O8

0 0 0 0 .012 .016 .020M = .5 .7 .9 .95

CD

.024 .028 .032 .036 •O4O

Figure 65, - SsXJETIII subsonic/transonic polars.

0o_,.-.4-)t_

%

00r.-

u,-

or..-

0

e--

0"10

C0

.r-.-

0")or.-

c"0

0

!

u:;

or-i,

\

\

/

_.J

Ce_

0

t.D

I..l.I

Z,<I---¢/)

_.1

X

I

Y

Z

//

_ ___Yn

Zn

(b) Modeling of the nacelle-wing interference problem.Shaded areas represent interference regions,

By

X

(a) Grid system for nacelle interference and winganalysis computations.

Figure 68. - Nacelle interference and wing analysis techniques.

ORIG//_AL PAG/_/SOF _X)_ Qu_J_rry

Figure 69. - Douqlas modeResearch Center

at:theNASA-AMES

OJ

0 e.=F:: .,,,...

,,_

_ O,=)0 4.0

0 0

_" II

m _

m m

• 7

I I

-J

! I

i°

ORIG/NAL PAGE/_

OF Poor qUAMTy

0

C_J

O00

n

!

!

("4

0

o_

COC) °

I!

I i I I I ! I _,-

! !

..J

0

0

!

O)

0E

CO 0

f,,.,.0

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Figure 74. - Correlation of maximum lift-to-drag ratio performancefor the Douglas model.

SSXJET

Sref : 89.65 m2 (965 ft 2)

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Figure 75. - SSXJET lift characteristics•

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MACH NUMBER

Figure 78. - SSXJET zero-lift drag components

SSXJET

M_= 2.2 NB= 16Sref : 89.65 m2 (965 ft 2)

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CDwave

CDwave

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NUMBER OF CUTS - NX

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Figure 79. - Convergence characteristics of the FAR-field wave dragmethod.

SSXJET

Two engines

Sre f = 89.65 m2 (965 ft 2)

.0028

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Figure 80. - Propulsion bleed and air conditioning drag.

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ACTUAL START CRUISES-_

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Figure 82. - SSXJET (L/D)ma x performance.

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CASE 2 /- ACTUAL

-----CASE 3 / START CRUISE.... CASE 4 CASE I

CASE 5

...... CASE 6

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MACH NUMBER

Figure 84. - (L/D)ma x performance for various SSXJET ! configurations.

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\\_ / WITH REDUCED FOREBODY CAMBER

%--

-_ TANDEM COCKPIT

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MACH NUMBER

Figure 86. - SSXJET II wave drag analysis.

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8.2

7.8

7.4

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6.6

6.2

5.8

_ TANDEM COCKPIT

ACTUAL START CRUISE POINTS

TANDEM_

SIDE-BY-SIDE

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l.O 1.2 1.4 l.6 1.8 2.0 2.2

MACH NUMBER

Figure 88. - Maximum L/D performance for the SSXJET II concepts.

SSXJETIII

GE21/JII-BIO ENGINESSref = 104.98 m2 (1130 ft 2)

•OOl 0

(-0

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MACH NUMBER

Figure 89. - Additive propulsion drag for the SSXJET III

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7.8

7.4

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6.6

6.2

5.8

5.4

ACTUAL START CRUISE L/D-_

5.0 I | i I _ .Ll.0 l.2 l.4 l .6 l.8 2.0 Z.2

MACH NUMBER

Figure 91. - SEXJET Ill maximum L/D performance.

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MACH NUMBER

I I I 1 I I

10455 11339 12558 13625 15240 17648

ALTITUDE, m

I I I I m ,,,I

34300 37200 41200 44700 50000 57900

ALTITUDE, ft

Figure 93. - "First cut" sonic boom estimates for the SSXJET.

10635 kg (23467 Ibm.)CRUISE AT OPTIMUM ALTITUDE

16442 m. (54700 ft.) OR CLIMB CEILING 7BEGIN CRUISE ALTITUDE /

434_ kg. (9576 Ibm.) /-

cLIMB ACCEL -_/

318 kg (700 Ibm)_I0 MIN, TAXl 7/

I .i, TAKEOFF//

L L 6191 kmF" 15407 kg

[r-=.

MAIN MISSION

19873 m. (65200 ft.)END CRUISE ALTITUDE

-- 102 kg. (225 Ibm.)

_f DESCENT. DECEL.

\ If8 kg. (260 Ibm)

MIN. TAXI

,_IP_,_ c_ _.m_./_ !TRIP FUEL (33968 Ibm.)

15526 kgT lBLOCK TIME AND FUEL

182 min. (34228 Ibm.)

NOTE: C.A.B. RANGE = TRIP RANGE MINUS TRAFFIC ALLOWANCEAS SPECIFIED FOR SUPERSONIC AIRCRAFT.

,8 at 8534 m. (28000 ft,)CRUISE AT BEST ALTITUDE 835 kg. (1841 Ibm.)

30 MIN. HOLD ATAND VELOCITY

1079 kg. (2378 Ibm.) / " _ (145_702fmt')_

_T_,_V / _s_v___ ,oo

No allowancey _- TO ALTERNATE AIRPORT "-_

MISSED APPROACH 1329 kg. (2930 Ibm.)

Figure 94. - SSXJET Mission profile

16612 m. (54500 ft.)

BEGIN CRUISE ALTITUDE

I0160 kg (22399 Ibm.)

CRUISE AT OPTIMUM ALTITUDE

OR CLIMB CEILING /4689 kg (10338 Ibm.) f

/318 kg (700 Ibm.) >z

I0 MIN. TAXI _ /

_ 6104 km

15266 kg

180 rain.

MAIN MISSION

TRIP RANGE

TRIP FUEL

]5384 kg

BLOCK TIME AND FUEL (33915 Ibm.)

19324 m. (63400 ft.)

END CRUISE ALTITUDE

99 kg (218 Ibm.)

DESCENT DECEL.

I18 kg (260 Ibm.)

MIN TAXI

(33655 Ibm.)

NOTE: C.A.B. = TRIP RANGE MINUS TRAFFIC ALLOWANCE

AS SPECIFIED FOR SUPERSONIC AIRCRAFT

Mach .8 at 9144 m (30000 ft.)

CRUISE AT BEST ALTITUDE

AND VELOCITY798 kg (]759 Ibm.)

i 30 MIN. HOLD AT4572 m

\1o69kg(23s6Ibm.) _ 7(15oooft.7% TRIP FUEL RESERVE "

No allowance __. _ TO ALTERNATE AIRPORT---_MT_'n APPROACHj _'D- TO"' ..... 1300 kg (2865 Ibm.)

Figure 95. - SSXJET I Mission profile

17160 m (56300 ft.)

BEGIN CRUISE ALTITUDE

4033 kg (8892 Ibm.) /

CLIMB ACCEL.--_ /

31a kg(70o1bin.) y

14328 kg

180 min.

9980 kg (21782 Ibm.)

CRUISE AT OPTIMUM ALTITUDE19782 m (64900 ft.)

END CRUISE ALTITUDE

OR CLIMB CEILING/

97 kg (213 Ibm.)

\f DESCENT DECEL.

MAIN MISSION

\\ I18 kg (260 Ibm)

_._ MIN TAXI

(3297 n.mi.)_ ]TRIP RANGE (31587 Ibm.)

TRIP FUEL 14446 kq--I _

BLOCK TIME AND FUEL (31847 Ibm.) -7

NOTE: C.A.B. = TRIP RANGE MINUS TRAFFIC ALLOWANCE

AS SPECIFIED FOR SUPERSONIC AIRCRAFT

Mach .8 at 9144 m (30000 ft.)

CRUISE AT BEST ALTITUDEAND VELOCITY 784 kg (1728 Ibm.)

30 MIN. HOLD AT4572 m1003 kg (2211 Ibm.) iF

n.mi.) ._No allowance / .,_ T482, _ER_O A I RPORTMISSED APPROACH --/ "- ILl AL "

277 kg (2815 Ibm.)

Figure 96. - SSXJET II Mission profile

17252m (56600 ft.)BEGINCRUISEALTITUDE

2671 kg. (8763 Ibm.) /

CLIMB ACCEL. -_ /

318kg.(70oIbm) yIOMIN. TAXI _ /

1 MIN. TAKEOFF//

6199 km.

F- 14321 kg.

9932 kq (21897 Ibm.)

CRUISE AT OPTIMUM ALTITUDEOR CLIMB CEILING 19903 m (65300 ft.)

END CRUISE ALTITUDE97 kg. (213 Ibm.)

\f DESCENT DECEL.

MAIN MISSION \

\118 kg. (260 Ibm.)

MIN TAXI

TRIP FUEL (31573 Ibm.)

14439 kg.

I-- 182 min. BLOCKTIME AND FUEL (31833 Ibm.)

C,A.B, = TRIP RANGE MINUS TRAFFIC ALLOWANCENOTE:

AS SPECIFIED FOR SUPERSONIC AIRCRAFT

Mach .8 at 9144 m (30000 ft.)

CRUISE AT PEST ALTITUDE 784 kg. (1729 Ibm.)

AND VELOCITY 7 30 MIN. HOLD ATI 4572 m

1002 kg (2210 Ibm.) / _ (15000 ft.)

MISSEDAPPROACHJ F'-- TO ALTERNATEAIRPORT-_1277 kg. (2516 Ibm.)

Figure 97. - SSXJET II Tandem mission profile

10547 kq (23252 Ibm)

CRUISE AT OPTIMUM ALTITUDE

17678 m (58000 ft.) OR CLIMB CEILING-7j

BEGIN CRUISE ALTITUDE D/

2608 kg (5750 Ibm) /#,

/318 kg (700 Ibm)

I0 MIN. TAXI

1 MIN. TAKE-OFF//

20117 m (66000 ft.)

END CRUISE ALTITUDE

113 kg (250 Ibm)

i ii Cii T (IIIEI bm)

MIN. TAXI

L 5375 km Trip Range (2902 n. mi.) _ !r 13586 kg Trip Fuel (29952 Ibm) _

_ 13704 kg -7

158 min. BLOCK TIME AND FUEL (30212 Ibm)

NOTE: C.A.B. RANGE = TRIP RANGE MINUS TRAFFIC ALLOWANCE

AS SPECIFIED FOR SUPERSONIC AIRCRAFT

Mach .9 at 12192 m (40000 ft.)

CRUISE AT BEST ALTITUDE I074 kg (2367 Ibm)

AND VELOCITY _/ 30 MIN. HOLD ATf 4572 m

1651 kg (3640 lbm)

Figure 98. - SSXJET Ill Mission Profile

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80 x 103

70

60

50

E4O

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I0

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20-

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Figure 99. - Mach-altitude climb schedule

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SSXJET DESIGN POINT

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25 30 35 .L-) 40 45

GROSS TAKE-OFF WEIGHT - kgI I I a I I I J

50 60 70 80 90 lO0 llO 120 x 103

GROSS TAKE-OFF WEIGHT - Ibm

SSXJET Ill CENTERLINE

WITH COANNULAR SUPPRESSION

I I

50 55 x 10 3

Figure I04. - Variation of maximum allowable EPNL with gross take-off

weight per 1969 version of FAR 36 (Ref. 29) at centerlinemeasurement station.

125 -

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0 SIDELINE

SSXJET IIISIDELINE

SSXjET IIISIDELINE WITH

COANNULAR SUPPRESSION

| I _ L l35 40 5 50 55 x 103

GROSS TAKE-OFF WEIGHT - kgI ,..i I I P

80 90 I00 llO 120 x 103

GROSS TAKE-OFF WEIGHT - Ibm

Figure 105. - Variation of maximum allowable EPNL with gross take-off

weight per 1969 version of FAR 36 (Ref. 29) at sidelinemeasurement station.

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CENTERLINE WITH

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I I I

40 45 50

GROSS TAKE-OFF WEIGHT - kgI I I I I I I

50 60 70 80 90 lO0 llO

GROSS TAKE-OFF WEIGHT - Ibm

f CENTERLINE

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I

55 X 103

!

120 x lO 3

Figure I06. - Variation of maximum allowable EPNL with gross take-off

weight per 1977 version of FAR 36 (Ref. 29) at centerlinemeasurement station.

125

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SSXJET

G SIDELINE

SSXJET III

SSXJET Ill

_ SIDELINE WITH

COANNULAR

SUPPRESSIONSIDELINE 4 ENGINES

SIDELINE 2 ENGINES

I I l I l

35 40 45 50 55 x ]03

GROSS TAKE-OFF WEIGHT - kgI I , I I i I

70 80 90 lO0 llO 120 x IO s

GROSS TAKE OFF WEIGHT - Ibm

Figure I07. - Variation of maximum allowable EPNL with gross take-off

weight per 1977 version of FAR 36 (Ref. 29) at sidelinemeasurement station,

STD. +IO°C DAY

15_=

!

i,i

-_ I0

F=-

5

20 x 102

8 x 102

FLAPS SET AT

.524 rad. (30 °)

6

0

I

0

CLIMB ANGLE = .161 rad. (9.25 °)

ALT = 329 m (1736 ft)

V = I17.5 msp (228 kt)

0 0

TART OF

TAKE-OFF ROLLTHRUST CUTBACK

FLAPS RETRACTED TO

.349 rad. (20°)IFTOFF

V = 91.4 raps

(177.6 kt)

RUNWAYCENTERLINE

___ef._ 29 MEASUREMENT£__POINTf 1

OBSTACLE_ft)lO.7m HEIGHT _"\ ,1981

I_- 'l \\\ 463 m (.25 n. mi.)m

6500 ft) I_ ,

6486 m (3.5 n. mi.)-t-_._RE _. 29 MESUREMENT POINT 2I

2 4 6 8 I0 12 x 103

DOWNRANGEDISTANCE FROM BRAKE RELEASE - m

I I I , I I I I I I I I I

1 2 3 4 5 6

DOWNRANGE DISTANCE FROM BRAKE RELEASE - n. mi.

*SIDELINE NOISE IS MEASURED WHERE NOISE LEVEL AFTER TAKE-OFF IS GREATEST

Figure 108. - Take-off profile for noise evaluation of SSXJET

STD. +IO°C DAY

20 x 102

8 x 102

FLAPS SET AT.524 rad. (30 °)

CLIMB ANGLE .142 rad. (8.12 ° )

ALT : 533 m (1750 ft)

V = 126.7 mps (246 kt)

6

START OFTAKE-OFF ROLL

LIFTOFF

V : 95.7 mps

(186 kt)

THRUST CUTBACKFLAPS RETRACTED TO

.349 rad. (20 °)

I

0

RUNWAYCENTERLINE

Ref.F/ _/- 29 MEASUREMENT POINT 1

BSTACLE HEIGHT P10.7 (35m

_'_ "-\ 463 m (.25 n mi.)1981 m o

(6500 ft) \6486 m (3.5 n. mi. _--_P'-_REF. 29 MEASUREMENT POINT 2

I I I I I J

2 4 6 8 I0 12 x 103

DOWNRANGEDISTANCE FROM BRAKE RELEASE - m

| I I I i I I _ I I I J1 2 3 5 6

DOWNRANGEDISTANCE FROM BRAKE RELEASE - n. mi.

*SIDELINE NOISE IS MEASURED WHERE NOISE LEVEL AFTER TAKE-OFF IS GREATEST

Figure 109. - Take-off profile for noise evaluation of SSXJET III.

1. Reix)rt No.

NASA TM 74055

4. Title and Subtitle

7 Author(s)

J 2. Government Accession No.

A PRELIMINARY STUDY OF THE PERFORMANCE ANDCHARACTERISTICS OF A SUPERSONIC EXECUTIVEAIRCRAFT

Vincent R. Mascitti

9 Performing Organization Name and Address

NASA Langley Research CenterHampton, Virginia

12 S_nsoring A_ncy Name and Address

National Aeronautics and Space AdministrationWashington, DC 20546

3 ReciDient's Catalog No.

5. Report Date

September 1977

6 Performing Organization Cede

31. lO0

8 Performing Organ,zatcon Repolt No.

10. Work Unit No.

11 Contract or Grant No.

13. Type of Report and Period Covered

Technical Memorandum

14 S_onsoring Agency Code

15 _pp_eme_tarv,o,_,Acknowledgment of analytic effort is as fol ows: Configuration Develop-

ment-E. E. Swanson and R. A. DeCosta; Stability and Control-G. L. Martin; Aerodynamics-

K. B. Walkley; Mass Characteristics and Mission Analysis-G. J. Espil; Propulsion-W. A.Lovell and Takeoff Performance and Noise-J. E. Russell(Vought Corporation; Hampton)

16 Abstract

A preliminary design study has been conducted to determine the impact of

advanced supersonic technologies on the performance and characteristics

of a supersonic executive aircraft. Four configurations with different

engine locations and wing/body blending were studied with an advanced non-

afterburning turbojet engine. One configuration incorporated an advancedGeneral Electric variable cycle engine and two-dimensional inlet with

internal ducting. An M 2.2 design Douglas scaled arrow-wing was used

throughout this study with Learjet 35 acconTnodations (eight passengers).

All four configurations with turbojet engines meet the performance goals of

5926 km (3200 n.mi.) range, 1981 meters (6500 feet) takeoff field length,

and 77 meters per second (150 knots) approach speed. The noise levels of

turbojet configurations studied are excessive. However, a turbojet with

mechanical suppressor was not studied. The variable cycle engine configu-ration is deficient in range by 555 km (300 n.mi.) but nearly meets sub-

sonic noise rules (FAR 36 1977 edition), if coannular noise relief is

assumed. All configurations are in the 33566 to 36287 kg (74,000 to 80,000

Ibm) takeoff gross weight class when incorporating current titanium manu-

facturing technology.

17 Key Words ISuggested by Authorls))

Supersonic CruiseSCAR

Aircraft Design

General Aviation

lB. Distribution Statement

Unclassified - Unlimited

19 S_urtty Cla,-,if, (of this repOrt)

Uncl ass ified I 20. SecurltvClassi|.(of this _)Unclassified

21. No of PaNs I 22. _ice"

191 I $7.50L

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