Acta Astronautlca Vol 4, pp 111-129 Pe rgamon Press 1977 Printed m Great Britain

Advanced supersonic cruise aircraft technology

H A L T. B A B E R , JR.t AND C O R N E L I U S D R I V E R ~ NASA Langley Research Center, Hampton, VA 23665, U S A

(Received 6 August 1976)

Abstract- -An early supersomc cruise vehicle conceptual study in 1969 Identified a number of configuration, stability and control, and performance problems Continuing research has provided soluhons to many of these problems, which were essentially aerodynamic in nature This article presents a multt-dtsclphne approach to the apphcatlon of the latest technology to supersomc cruise aircraft concept definition, identifies current problem areas and cites developing technology, in several dtsclphnes, which are prospective solutions to these problems Approach, status and potential for improvement are addressed through consideration of a Langley Research Center (LRC) concepted supersonic cruise configuration (based on technology developed since 1969) and several parametric variations As part of the developing technology, a recently developed aircraft sizing and performance computer program now makes possible rapid analysis of the effects of configuration resizing such as changes m wing loading and thrust-to-weight ratio This program was used to determine allowable wing loading and takeoff gross weight sensitivity to structural weight reduction Structural weight reduction as large as 15%, or more, appears highly probable m hght of recent technology advances with fiber reinforced composites or other advanced structural techniques Based on scale model tests of coannular nozzles, noise suppression on the order of that required for the configurations considered here can be achieved

Introduction THE LANGLEY Research Center, since U.S. Congressional approval in 1972, has been actively engaged in, and contractually supporting, work in advanced supersonic technology for potential application to future U.S. transport aircraft. One of the traditional methods of assessing these improvements in technology is through the use of conceptual integrated system studies. These types of studies result in hypothetical reference airplanes which allow assessment of the tech- nology improvements in terms of increases in range/payload or reduction in gross weight. At the conclusion of NASA sponsored studies in 1969 (The Boeing Company, 1969) the reference configuration was a 340,194kg (750,000 Ibm) aircraft with a payload of 234 passengers and a range of 6760 km (3650 n.mi.). This configuration required a fold-out canard in the low-speed flight regime for trim during takeoff and landing with the horizontal tail providing pitch control. However, this high-lift configuration had a pitch-up problem at relatively high lift coefficient. Because of non-linear pitching moments in the low-speed regime a hard SAS (Stability Augmentation System) was not feasible.

In the interim, since 1969, results (unpublished) from extensive Langley Research Center low-speed wing tunnel tests have led to significant improve- ments in longitudinal stability at high angle-of-attack for the takeoff and landing modes. This has been achieved by careful attention to wing planform, leading-

tAerospace Engineer, Vehicle Integration Branch ~tHead, Vehicle Integration Branch. Aeronautical Systems Division

111

112 H. T Baber, Jr. and C. Driver

edge radius, leading-edge high-lift devices and trailing-edge flap location, size and deflection. Improvement in stability allowed adoption of a hard SAS and elimination of the canard, thereby reducing drag and weight. These results along with advancements in analytical techniques have, therefore, indicated potential improvements in supersonic aerodynamic technology.

This new technology was utilized to evaluate wing thickness variations, nacelle contouring for minimum drag at cruise and the use of the horizontal tail in a lifting attitude during climb and cruise to increase lift-to-drag ratio for performance improvement. The current baseline configuration, designated AST- 100, is the outgrowth of this activity. AST refers to advanced supersonic technology. Criteria for the concept definition study were as follows:

• Five abreast seating of 292 passengers, all tourist class with seat pitch of 0.864m (34 in.);

• Standard day cruise at Mach number (M)= 2.7; • Range of 7408 km (4000 n.mi.) on standard day +8°C at M = 2.62; • Engine size based on noise considerations (Dept. of Transportation, 1974),

transonic acceleration (with hot day thrust margin of 1.2), or cruise, whichever is critical;

• Land on existing runways with tire footprint no greater than that of DC-8-50.

• Stability and control (abbreviated list--see Baber and Swanson, 1976, for complete criteria):

No significant pitch-up in the takeoff or landing modes; Cruise static margin at 1.0g of at least 3% of the reference chord to offset

reduction in stability due to flexibility at the required 2.5g pitch-up maneuver; Cruise directional stability such that the yawing moment derivative, C~ B, is

equal to or greater than zero for the 2.5g maneuver. Baber and Swanson (1976) studied the following technical areas: configura-

tion development, aerodynamics (low and high speed), stability and control, mass characteristics, noise, sonic boom, mission analysis (performance) and direct operating cost. This article is based on the Baber and Swanson study but the scope is limited, primarily, to noise and performance analyses. Analytically obtained noise and performance results are presented for the current baseline concept and for several configurations quite similar to the baseline. These latter configurations differ from the baseline only by variation of one or more parameters such as thrust-to-weight ratio, wing loading and takeoff gross weight.

AST-IO0 advanced supersonic cruise vehicle Description

The AST-100 wing profile is the result of several iterations and was selected as having characteristics nearest to optimum with achievable geometry. For example, the chordwise location of maximum thickness is as far rearward as practical and, among the wings considered by Baber and Swanson (1976), it has the lowest volume and hence the lowest wave drag. A thickness map of the selected wing is shown in Fig. 1.

After locating the wing to satisfy balance requirements, sizing the empennage

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accordingly and contouring nacelles for minimum drag at cruise, the fuselage was recontoured for minimum drag.

The design features and fabrication techniques for major components were assumed to be as follows:

Wing and aerodynamic surfaces--titanium/core sandwich panels; • Fuselage--titanium skin/stringer/frame construction.

The configuration general arrangement is shown in Fig. 2. Seating is five abreast, all tourist class, to accommodate passengers in accord with the criteria. Basic mission fuel is contained in integral fuel tanks in the wing. A 15142 1 (4000 gal.) reserve fuel tank for extended range and/or center-of-gravity control is located in the aft fuselage. The main landing gear is a two-strut arrangement with twelve wheels per strut. Since detailed design of the main gear was beyond the scope of the study, volume for clearance was provided based on wheel diameter and strut length only. The four engines are single spool, variable geometry turbine, non-afterburning turbojets. They are rated at 323 kg/sec (712 Ibm/sec) cbrrected compressor airflow (uninstalled) each.

Aerodynamic characteristics Low speed. Low speed aerodynamic characteristics were derived primarily

from Coe et al. (1975), which is an example of recent high-lift configuration

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research applicable to a supersonic cruise vehicle in the takeoff and landing modes. Adjustment of this data was made to account for skin friction differences due to the difference between model and full-scale Reynolds numbers. In addition, corrections were made for differences in leading edge suction between model test and full-scale flight conditions by the use of an equation (modified) from Henderson (1966). Supplementary drag data attributable to air conditioning and propulsion were taken from a Boeing Commercial Airplane Co. report (1974), and that for the landing gear was from The Boeing Company (1969). These results were summed and plotted as drag polars for the three low-speed flight modes, namely, takeoff and initial climb, climbout, and approach. Lift-to- drag (L/D) ratios as obtained from these polars are shown in Fig. 3. Technology advancements, such as low-speed wind tunnel test data for an arrow wing in lieu of empirical estimates of lift and drag and the availability of experimental leading-edge suction data, which made possible the application of leading edge suction scaring effects, have resulted in higher confidence than heretofore in the prediction of high-lift characteristics for the supersonic cruise vehicle.

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Advanced supersonic cruise aircraft technology 117

Moreover, lift-to-drag ratios for the three low-speed flight modes are slightly higher (0.26 to 0.82) than could have been expected prior to 1975.

High speed. Overall high-speed drag polars for the AST-100 were developed by combining zero-lift drag increments (including air conditioning and propulsion bleed drag) with semi-empirical estimates of the horizontal tail contribution to lift and drag due to trim incidence, and characteristic polars from a previous analysis (LTV Aerospace Corp., HTC, 1973). Specifics on the methods applied can be found in Baber and Swanson (1976).

From such overall drag polars, maximum lift-to-drag ratios were graphically determined and are shown in Fig. 4 as a function of Mach number.

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Noise and chemical emission In accord with Department of Transportation (1974), two flight conditions for

the AST-100 were analyzed for noise. They were takeoff and approach, with both centerline and sideline measurement stations considered for the former and centerline only for the latter. The required location of these measurement points relative to the runway and flight path is illustrated in Fig. 5. The environmental conditions for the analysis, as prescribed by Department of Transportation (1974), were temperature at standard day +10°C, and relative humidity at 70%. Hence, engine performance, speed of sound, and atmospheric absorptivity for noise analysis were based on these conditions. However, it should be recalled

118 H.T. Baber, Jr. and C. Driver

APPROACH TAKEOFF

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wcNOTE SIDELINE NOISE IS MEASURED WHERE NOISE LEVEL AFTER LIFTOFF IS GREATEST

Fig. 5 Noise measurement |ocations for approach and takeoff

that aircraft performance was based on standard day +8°C as per the study criteria.

Prior to 1973, noise analysis was dependent upon a 1965 source for basic inputs such as the manner of variation of overall sound pressure level (OASP L) with jet relative velocity; of density exponent with jet relative velocity; and of the directivity correction with frequency and jet relative velocity. Since 1965, noise research has resulted in an improvement in knowledge in the manner of variation of these parameters, which is essential to the mathematical model representation of aircraft noise. With OASPL as an example, current, more accurate estimates give AST-100 takeoff noise as 7.5 decibels (dB) lower than would have been predicted before 1973.

Takeoff noise analysis. Takeoff noise level evaluation consisted of an engine sizing parametric study and a final noise analysis compared to the criteria of Department of Transportation (1974). The effects of trailing edge flap angle and engine power setting on takeoff noise level at the measurement stations shown in Fig. 5 were included in the study.

The effect of takeoff distance on effective perceived noise level (EPNL) at the sideline measurement point was also determined as a function of flap deflection and power setting. For a given power setting (jet exit velocity), sideline noise is slightly dependent upon takeoff distance and flap deflection and, conversely, more strongly dependent upon engine power level.

Installed thrust-to-weight ratio (TIW) varies for each power setting (jet exit velocity). By using the appropriate range of T[W values, the variation in noise level with jet exit velocity was obtained and then related to uninstalled TIW's for preliminary engine sizing.

Required noise suppression will be a minimum for the condition of equal noise at the centerline and sideline measurement points. For the AST-100, this

Advanced supersonic cruise aircra[t technology 119

occurs at a jet relative velocity of 752.9m/sec (2470ft/sec) for which the centerline and sideline noise is l l4.4dB. Department of Transportation (1974) limits the EPNL at the two takeoff measurement points and the measurement point for approach to 108dB. However, noise tradeoffs are allowed between the three measurement points with the provision that no more than 2dB can be applied to noise exceedence at any point and that the sum of exceedence cannot be greater than 3dB. The approach noise, to be discussed subsequently, is less than 108dB. Since Department of Transportation (1974) permits 3dB of the difference between the allowable noise level and the actual approach noise level to offset the exceedence at the two takeoff noise measurement points, the limits at these points can be raised to 109.SdB. Required suppression is therefore 4.9dB.

Contour plots, for the AST-100, of EPNL (unsuppressed engine) for 108 and llSdB are presented in Fig. 6.

Approach noise analysis. For the AST-100 in the approach mode, noise level was evaluated for the measurement point defined in Department of Tran- sportation (1974), and illustrated in Fig. 5. For the approach, trailing edge flaps were deflected to 20 degrees. Lift coefficient during 3-degree approach is 0.55. The normal landing weight of 205,930 kg (454,000 Ibm) was used in this analysis. At these conditions, the approach speed is 81.75 m/sec (158.8 knots) and the LID is 5.9. With the aircraft in these approach conditions, the EPNL at the applicable measurement station was calculated to be 103.1dB.

Airframe noise analysis. The airframe overall sound pressure level (OASPL)

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DOWNRANGE DISTANCE FROM BRA~E RELEASE n ml

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120 H.T. Babcr, Jr and C. Driver

was computed by use of the following equation:

V3~WO~ ] O A S P L = 10 log,o R. 6-~6-V~-eo6j + 41.3

where V = aircraft speed in m/sec W = aircraft weight in kg R = aircraft altitude in m S = wing area in m 2

A R = wing aspect ratio 41.3 = constant for correlation of empirical and experimental data.

The airframe O A S P L value predicted by this equation is applicable to the runway centerline directly under the aircraft. The O A S P L values so determined are 108 and 107dB at the takeoff and approach measurement points, respectively. Since O A S P L is not time dependent , these values are the maximum that an observer at either measurement point would perceive. Because E P N L is a time-related function, the Department of Transportat ion (1974) EP N L limitation is not directly applicable to airframe OASPL. However , the E P N L will be less than 108dB if the O A S P L is equal to or less than 108dB.

Chemical emissions. The AST-100 engine is a dry turbojet with a variable area turbine to keep the airflow at a relatively high level (for reason of noise) following thrust "cu tback" after takeoff. This engine, which is a scaled-down relative of the General Electric Company's GE-4 turbojet of 1971, would have oxides of nitrogen emissions, at cruise, on the order of 18 g/kg of fuel consumed. For.a typical mission, during the cruise phase only, this would be approximately 1520 kg (3350 Ibm). Meanwhile, in the five year interim, NASA's Lewis Research Center and engine manufacturers have developed jet engine combustor tech- nology to decrease such pollutants to an emission index of 7.8-8.8 g/kg of fuel, based on scale model test data.

Mission analysis

Mission analysis objective for the AST-100 was a design range of 8408 km (4000 n.mi.) and 292 passengers plus baggage, which translates into a payload of 27,682kg (61,028Ibm). Although the AST-100 is sized for M = 2.7 cruise at standard conditions, it is required that it achieve the range objective on a hot (standard +8°C) day. On the basis of equal stagnation temperature, a standard +8°C day Mach number of 2.62 corresponds to M = 2.7 at standard conditions and hence was selected as the cruise Mach number for mission analysis in order not to exceed M = 2.7 stagnation temperature. The desired mission profile is presented in Fig. 7. Fuel reserves allowance from F A Rt 121.648 SST Fuel Requirements (tentative standard proposed by the FAA$) was modified for a change in holding altitude from 457m (1500ft) to 4572m (15,000 ft).

tFederal Aviation Regulation ~tFederal AviaUon Administratton

Advanced supersonic cruise aircraft technology 121

CRUISE AT OPTIMUM ALTITUDE OR CLIMB CEILING-~

START CRUISE ALT END CRUISE ALT 18698 rn 21031 m

(62000 ft ) . 207995 ,.~'-~(69000~ f t ) / 292300 kg r,,.j \

10 MIN TAXI + ~ 1 MIN TAKEOFF~ / DESIGN MISSION

(37~o6J~l~CJrn) / / DESIGN RANGE 7408 km(4000n m,) . . . . . \ ~-5 MIN TAXI • ( 321325 kg 206385 kg %, ,I

(708400 ibm) (455000 Ibm)/ I

t - - TRIP RANGE 73L,9 kmt3966 n ml) ~'l /

BLOCK FUEL 120464 I~g(265578 Ibm)

NOTE C A B RANGE=TRIP RANGE MINUS TRAFFIC ALLOWANCE AS SPECIFIED FOR SUPERSONIC AIRCRAFT

Fig. 7. Design mission performance profile

Sizing in terms of engine size and overall aircraft weight was performed. This sizing was accompl ished through a parametric study, which determined the effect of TIW ratios on range for various takeoff gross weights.

Plots of takeoff gross weight (TOGW) as a funct ion of T/W for the design range and range as a funct ion of T/W for various TOGW's were generated. For an uninstalled T/W = 0.41 and a design range of 7408km (4000 n.mi.), a T O G W of 325,679kg (718,000 Ibm) was obtained from these plots. A weight summary, which is a coarse breakdown by major e lements , is presented in Table 2.

Table 2 AST-100 weight summary

KILOGRAMS POUNDS

STRUCTURE 87576 193073

PROPULSION 27016 59561

SYSTEMS 2706~ 59665

WEIGHT EMPTY 141656 312299

OPERATING ITEMS 779t 17177 OPERATING WEIGHT 149448 329z,76

PAYLOAD 27682 61028

ZERO FUEL WEIGHT 177130 390504

MISSION FUEL 148550 327496

TAKE-OFF GROSS WEIGHT 3256?9 718000

The N A S A Mission Analysis Computer Program (PAB-2011) was then used for a final miss ion performance evaluation of the AST-100. Significant results

122 H T Baber, Jr and C. Driver

are:

(1) Start of cruise L/D (2) Start of cruise altitude (3) Required trip fuel (4) Range

----8.91 --18,898 m (62,000 ft) - - I 19,295 kg (263,000 Ibm) --7,349 km (3968 n.mi.)

Extensive mission segment data are presented in Table 3. A simplified flight profile representation of mission performance is shown in Fig. 7 along with the design mission definition.

Computed performance of the AST-100 indicates a range that is 59.3km (32 n.mi.) short of the objective of 7408 km (4000 n.mi.) or 0.8% less. This small difference was considered negligible and further iterations were not made.

Supersonic technology development benefits Although recent supersonic technology development benefits are too nu-

merous to detail here, they can perhaps be most easily understood and ap- preciated by examination of a short, summary type table (Table 4). This table makes possible a comparison of "where we were" in 1969 with "where we are" in 1976, based on today's technology. The data show that more passengers could now be transported a greater distance with a smaller (lower TOGW) aircraft. This kind of performance gain is largely attributable to the improvement in aerodynamics. This performance gain will, of course, significantly improve the operating economics of such transports. In addition, air pollution at cruise altitude due to jet engine operation can be reduced by about 55% based on estimates using scale model test data as inputs.

Advanced supersonic technology applied to AST-100 resizing

Background and description A Langley developed aircraft sizing and performance computer program

(Fetterman, 1976) has been used to rapidly analyze the effects of configuration resizing such as changes in wing loading, W/S, and thrust-to-weight ratio, T/IV. A typical sizing plot for an AST-100 type aircraft is shown in Fig. 8. As a point of reference, the AST was sized at T / W = 0.37 (installed, std. day +8°C) and W/S = 351 kg/m 2 (721bm/ft2), based on wing reference area, S. The extent to which W / S can be increased to improve range is limited by takeoff field length performance which is dependent upon low-speed lift capability, and by the noise constraint, bounded by takeoff speed and which, in turn, is inversely related to lift-to-drag ratio. Improved low-speed data (Coe et al., 1975) for supersonic cruise aircraft along with continued development of coannular nozzles for noise relief should relax these takeoff related constraints.

With the AST-100 configuration as the baseline, the sizing and performance program was applied to determine the maximum allowable wing loading to meet or exceed the range criteria of 7408 km (4000 n.mi.) within the constraints of AST-100 TOGW, fuel limit, takeoff speed and field length. Results are plotted in Fig. 8. It can be seen that 415 kg/m 2 (85 lbm/ft 2) is the maximum possible wing

Advanced supersonic cruise aircraft technology 123

Table 3 Mission performance

MISSION Deslgn Supersonlc Crulse Math 2 62

MODEL NO AST-IO0

#IRCP~FT CHARACTERISTICS

Takeoff gross welght kg (Ibm) 325679 (718000) Operatlnq weight empty - kg (Ibm) 149448 (329476) Payload-No Passengers - 292

Cargo - 0 Total Weight kg (Ibm) 27682 (61028)

Wing area - reference m2 ( f t2) 926 (9969) - actual m2 ( f t2) I022 (I0996)

S L stat lc Instal led thrust per englne (std day +8 C), N ( Ib f ) 293485 (65978)

I n l t l a l instal led thrust to weight ra t lo 37 I n l t l a l wlng loadlng - reference, kg/m 2 (Ibm/ft2) 353 (72 O)

actual, kq/m2 (Ibm/ft 2) 319 (65 3)

Design Mission

Takeoff

Start Cllmb

Start Crulse

End Cruise

End Descent

Taxl- ln

Block Fuel and Time

Trlp Range

NOTES 1

OPERATING & FUEL G RANGE & TIME WEIGHTS, kg (Ibm) kg (Ibm) kln (n m ) mlnute

325679 (718000) 4354 (9600) 0 II

321325 (708400) 28925 (63769) 624 (337) 22

292400 (644631) 84405 (186081) 6354 (3431) 134

207995 (458550) 1610 (3550) 370 (200) 20

206385 (455000) I169 (2578) 0 5

120463 (265578)

7349 (3968)

Taxl- ln fuel taken out of reserves at destlnatlon

192

C A B range correspondlng to block tlme and fuel equals t r lp range mlnus t r a f f l c allowances as wl l l be sDeclfled for supersonic a l r c ra f t

Model No AST-IO0

Reserve Fuel Breakdown, kg (Ibm)

1 7% Trlp Fuel 8351 (18410)

2 Missed Approach 4037 18900)

3 482 km (260 n m ) to al ternate alrport 10482 (23108)

4 30 inln holdlng at 457 in (15,000 f t ) 6383 (I~0731

Total Reserve 29253 (64491)

I n l t l a l Crulse Condltlons

L i f t Coeff lc lent 0980

Drag Coeff lc lent DllO

Lif t /Drag 8 91

TSFC, kg/hr/N ( Ibm/hr/ Ibf) 0 138 (I 355)

A l t l tude, m(ft) 18898 (62000)

124 H.T. Baber, Jr and C. Driver

Table 4. Benefits of supersomc technology development

SST-969-336C (SCAT-15F) AST*

1969 1976

Takeoff gross weight, kg(Ibm) 340,194(750,000) 325,679(718,000)

Takeoff maxlmum noise (meas pts ), EPNdB 125 If4

Passengers 234 292

Payload, kg(Ibm) 22,180(48,906) 27,682(61,028)

Cruise Mach number 2 7 (Std day) 2 62 (Std day ÷8°C)

Crulse aerodynamlc ef f lc lency, L/D ~ ~4 8 91

Cru]se NO x emlsslon, g/kg fuel (estlmated) 18 1 8-8 8

Range, km (n m] ) 6,760(3,650) 7,348(3,968)

* Advanced Supersonic TechnoIoql

%

llao,Jroach, m/s(kts) 92.6(1J0) 3Z 'J /

Pan(je, k , , , ( I ) . . . . // // 67(]u03)---- ~

73(4200)-- _

3 8519(-1600 }-- ~__~_ /' ~ E / ~ - : z ~ r -/t

Vtake0f f , ~ / s ( k t s ) = ] 3 2 9(200)

s(170) /

/

i -

53,3 ~ '30 .00 ] 50

82 3(160) /7 L(150)

/ /

i/ Thru!

J . / J

~ F l e l d L engtn,

300 250

Takeoff 'd/S, kqlrn 2

Margln ( T / D - I ~

]O,Crulse

25,C11mb

,~ Ft):-°O0(10300)

11 n 100 Ol] ~0 70 6 ~ 50

Takeoff [J/S, lb, i,'f~

Fig 8. Results of General Electric Co turbofan scale model jet noise tests.

loading. Further it should be noted that the minimum installed T! W to meet the takeoff field length constraint is 0.31, which is 0.06 lower than that of the AST-100. With this new wing loading, a configuration, designated AST-102, has been defined. The AST-102, for the same TOGW as the AST-100, has a reference wing area of 784.76 m s (8447 ft2), which is slightly smaller than that of the AST-100 due to the higher wing loading. This corresponds to a gross wing area of 856.6 m s (9317 ft 2) for the AST-102.

Changes in the W/S and T! W necessitated rework of several design areas as deviations from the AST-100 baseline before detailed analysis of the AST-102 could proceed. These were:

• Landing gear length;

Advanced supersomc cruise aircraft technology 125

• Local wing thickness changes to accommodate landing gear; • Low speed lift curves and drag polars; • High speed drag polars (particularly skin friction and wave drag con-

tribution); • Horizontal and vertical tail resize; • Engine weight and nacelle geometry scale down.

Noise compar i son Engine and airframe noise levels were predicted for the AST-102 concept in

the takeoff and approach modes of flight, by the use of the same computational procedures as for the AST-100 configuration.

Table 5 presents the calculated noise for the AST-102 in the high lift condition applicable to takeoff and landing. For comparison, AST-100 results at like conditions are also presented in this table. It is apparent from these results that with the 3dB of noise tradeoff available from the approach, as in the case of the AST-100, the required suppression for the AST-102 is 12.4dB to comply with the Department of Transportat ion (1974) limitations of 108dB.

Table 5. Predicted noise levels at the FAR* 36 measurement points

Runway Alr f rame Cen te r l l ne S~del lne Nolse EPNL - dB EPNL - dB OASPL - dB

Takeoff I14 4 I14 4 108

Approach 103 1 107 y

Takeo f f 122 1 121 7 108 ~1 I

Approach 95 7 107 I

AST-IO0

AST-I02

* Federal Av~at lon Regulat lon

Due to recent technology development in the area of coannular nozzles, noise suppression on the order of that required for the AST-102 is believed to be achievable. This is supported by scale model test data from coannular nozzles. Typical data from such tests are shown in Fig. 9. These data indicate that a nozzle designed to exhaust a cool, low-velocity stream surrounded by a relatively hot, high-velocity concentric jet stream has considerable potential for lower noise level than the conventional, conical nozzle. Quantitatively, it is 10-12 E P N d B less in the applicable range (see AST-100 Noise and Chemical Emissions discussion) of takeoff jet exit velocity for supersonic cruise vehicles.

For a better perspective of the impact of noise reduction due to application of coannular nozzle technology one should refer back to Fig. 6. In this figure the dashed curve, which is a 104.4dB contour, represents a 10dB reduction in AST-100 noise if the conventional conical nozzles be replaced with coannular

126 H.T. Baber, Jr. and C Driver

T w

~ I0 dB ~_ co

o:"

t BASE /

/ VFA~, I " 1,5 VCORE,

'l 5 6 7 ~ 9 x 10 2

BYPASS JET VELOCITY, ,, 's k I L I __J_

1s -1'~ - - ~ ~ ~ 1 0 2 P, YPASS JET VELOCITY, f t ' s

F~g 9 Supersomc airplane parametric s~zmg

nozzles. The obvious message here is that the regulatory noise limit can be confined to a significantly smaller piece of real estate around the airport.

In view of accompl ishments to date and on-going activity in this technical discipline, it would appear that coannular nozzle technology for noise relief will be available by 1980, with a possible service entry date of 1984.

AST-I02 Performance The AST-102 operat ing weight was less, by 9917 kg (21,864 Ibm), because of

reduced wing area and lower T/W (hence lighter engines) compared to the AST-100. This weight change would suggest further aircraft resizing. However , since the initial resizing design philosophy was to maintain the same TOGW, further resizing was temporari ly held in abeyance. This est imated reduction in operating weight was replaced by an equal weight of fuel. Volume is available in the aft fuselage tank for fuel in excess of the capaci ty of the wing tanks.

An in-depth mission analysis was per formed for the AST-102 at the same takeoff constraint, flight environment , and weight conditions as for the AST-100, with the exception of the tradeoff of some operating weight for additional fuel as just described. This analysis indicated that the AST-102 at a T O G W of 325,679 kg (718,000 Ibm) could t ransport 292 passengers and baggage a distance on the order of 8260 km (4460 n.mi.) at M = 2.62 on a standard +8°C day. This point is shown in Fig. 10, as the AST-102 baseline, on the curve for zero percent structural weight improvement .

In Fig. 10 the curves of range as a function of TOGW, for various percentage changes in structural weight, were generated by use of the aircraft sizing and per formance computer program, mentioned previously. This was done to inves- tigate AST-102 range and T O G W sensitivities to structural weight reduction. These curves can be used in either of two ways. For example , the plot can be

Advanced supersomc cruise aircraft technology 127

50:

45

~ 40 = w

= " 3 5

30

25

:103

E

9 5xlO 3

8 5

7 5

65

/

s s / /

4 5 220

FUEL LIMIT ~

/ STRUCTURAL WEIGHT , ~ REDUCTION' 1 5 ~ / /

240 260 280

1 - f

/ / /

[

AST-I02 BASELINE 2 0 W,/SRE F = 415 kg/m (85 Ibm/ft 2)

300 320 340 360 x lO 3

TAKEOFF GROSS WEIGHT, kg I I i I I |

500 550 600 650 700 750 x 103

TAKEOFF GROSS WEIGHT, Ibm

Fig. 10 Range and takeoff gross we ight sens i t iv i ty to structural we ight imp rovemen t .

entered at constant range, say 7408 km (4000 n.mi.) to determine the required TOGW for a given structural weight expressed as a percentage change from the baseline structural weight. Conversely, the plot can be entered at a fixed TOGW, for example 325,679 kg (718,000 Ibm), to determine range for various percentage changes in structural weight from the baseline.

Structural weight reduction as large as 15% is projected here. This much improvement, or more, appears highly probable in light of recent technology advances with fiber reinforced composites (see Boeing Commercial Airplane Co., 1974) and other advanced structural techniques. Such structural weight reduction, for constant range, would result in takeoff gross weight reduction of about 10% (from Fig. 10), with attendant improvement in noise and operating efficiency. The potential for overall structural weight improvement is illustrated by the panel samples shown in Fig. 11.

Based on current status and expected level of development, 1986 is the projected technology availability date for fiber reinforced composites. Experi- mental determination of composite material high temperature/strength charac- teristics and high temperature-long term exposure (with cycling) physical properties must be accomplished in the interim to provide design data by that date.

Concluding remarks

Technology advances since cessation of the U.S.A. SST development effort give promise of remedying deficiencies which rendered that 1971 concept environmentally unacceptable and economically unwise.

128 H T Baber, Jr and C Driver

Fig. 11 Composite panels for supersomc aircraft

A conceptual s tudy using state of the art supersonic technology, which included recently developed "methods of analysis and aerodynamic data f rom wind tunnel tests of supersonic cruise configuration models, has defined an aircraft which would be capable of transporting 292 passengers plus baggage approximate ly 7408 km (4000 n.ml.) on a standard day +8°C at M = 2.62. Within current technology this configuration (designated AST-100) is believed to be a viable concept which approaches environmental acceptabil i ty and economic profitability. Jet engine technology deve lopment since 1969 gives promise of lower (approx. 10dB) unsuppressed takeoff noise level and an emission index that is lower by about 55%.

These advances plus potential engine noise suppression by use of coannular nozzles, and structural weight fraction reduction by application of fiber rein- forced composi tes are reasons for opt imism in future supersonic cruise vehicle design.

References

Baber, Hal T, Jr and Swanson, E E (1976) Advanced supersomc technology concept AST-IO0 charactenstws developed In a basehne-update study. NASA TMX-72815

Boeing Commercial Airplane Company (1976) Studies of the impact of advanced technologies applied to supersonic transport aircra[t, Task III concept refinement and engine coordination I)6- 22558.

Coe, P. L., Jr, McLemore, H. Clyde and Shivers, J P (1975) Effects of upper-surface blowing and thrust vectoring on low-speed aerodynamic charactenstws o[ a large-scale supersonic transport model NASA TMX-72792

Advanced supersomc crutse atrcra[t technology 129

Department of Transportation (1974) Federal Aviation Administration no~se standards" aircraft type and airworthiness certification Federal Aviation Regulations, Part 36. Appendix C, pp C36 I - C.36.5.

Fet.terman, D. E., Jr. (1976) Preliminary sizing and performance evaluation of supersonic crmse aircraft. NASA TMX-73936.

Henderson, W P. (1966) Studies of factors a~ectmg drag due to hft at subsonic speed. NASA TN-D 3584

LTV Aerospace Corporation, HTC (1973) Advanced supersomc technology concept study, reference characteristics NASA CR-132374

The Boeing Company (1969) Mach 2 7 fixed wing SST model 969-336C (SCAT-15F). Document No DGA-11666-1