NASA/TM-1999_209534

Nacelle Integration To Reduce the Sonic

Boom of Aircraft Designed To Cruise at

Supersonic Speeds

Robert J. Mack

Langley Research Center, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research Center

Hampton, Virginia 23681-2199

December 1999

https://ntrs.nasa.gov/search.jsp?R=20000014072 2018-05-14T14:09:50+00:00Z
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Nacelle Integration To Reduce The Sonic Boom

Of Supersonic-Cruise Aircraft

Robert J. Mack

Langley Research Center

Hampton, Virginia

SUMMARY

An empirical method for integrating the engine nacelles on a wing-fuselage-fin(s)

configuration has been described. This method is based on Whitham theory and Seebass and

George sonic-boom minimization theory. With it, both reduced-sonic-boom as well as high

aerodynamic efficiency methods can be applied to the conceptual design of a supersonic-cruise

aircraft. Two high-speed civil transport concepts were used as examples to illustrate the

application of this engine-nacelle integration methodology: (1) a concept with engine nacelles

mounted on the aft-fuselage, the HSCT-10B; and (2) a concept with engine nacelles mounted

under an extended-wing center section, the HSCT-11E. In both cases, the key to a significant

reduction in the sonic-boom contribution from the engine nacelles was to use the F-function shape

of the concept as a guide to move the nacelles further aft on the configuration.

INTRODUCTION

The conventional location for engine nacelles on a supersonic-cruise aircraft is under the

wings. This is the arrangement seen on the United States' XB-70, the English-French Concorde,

and the Russian Tu- 144, references 1 to 3. It is desired because the engine inlets are in a favorable

flow field under the wing where the shock losses from the nacelles can be recovered through

nacelle-wing interference lift. Most conceptual high-speed civil transport (HSCT) aircraft studied

in subsequent years followed this engines-under-the-wing practice. References 4 to 10 document

but a few of the configurations that have appeared during the time period when the Supersonic

Transport (SST) Program and the follow-on Supersonic Cruise Aircraft Research (SCAR)

Program funded studies of design, performance, maintenance, and marketing problems.

When the engine-nacelle location was changed, it was usually for a specific reason. One such

reason was the reduction of the aircraft's sonic-boom on the ground. Sonic-boom minimization

theory, reference 11, and sonic-boom reduction methods, references 12 to 14, showed that ground

overpressure noise could be reduced by careful shaping of the aircraft's geometry. These methods,

however, assumed that the aircraft's components had areas that were smooth and continuous. It

was found that engine nacelles did not meet this criteria, so they had to be treated differently. In

this paper, the reasons for moving the engine nacelles aft to reduce sonic-boom are analyzed and
discussed.Theconceptualaircraftin thesediscussionswerethoseinstrumentalin discoveringand

isolatingnacelle-integrationeffectsonsonicboom,aswell assubsequentconceptualaircraft

whoseunconventionalengine-nacellepositionswouldenablethemto cruisesupersonicallywith

reducedgroundoverpressure.

SYMBOLS

CL

CL,CRUISE

F(y)

h

M

W

lift coefficient

lift coefficient at beginning of cruise

Whitham F-function of variable y

altitude or separation distance, fl

Mach number

aircraft weight, Ib

longitudinal distance along model signature, in

effective length along the longitudinal direction of the aircraft or concept

BACKGROUND

The impact of nacelles on the shock systems from conceptual aircraft was first observed

during wind-tunnel measurements of pressure signatures from the Mach 2 and the Mach 3 theoryvalidation wind-tunnel models of reference 15. These were 1:600 scale models of conceptual

aircraft designed with the methodology documented in references 16 and 17 as well as

methodology reflected in the wind-tunnel models and pressure-signature measurements described

in reference 18. A three-view sketch of the Mach 2 concept is shown in figure 1.

c

c

.....---q__

Figure 1. Mach 2 low-boom theory-validation concept.

2
Duringthewind-tunneltests,thismodelgeneratedapressuresignaturewhich hadanunexpected

strongshockbetweenthenoseandtail shocksasseenin figure2.

0.080 v

I'i.03 _ ========================

- .Q30

_!iiiiii!iiiiiiiii

_j_j_..:.-ij..j_j

B.o I_.o 24.0 _..o

X._in

Figure 2. Pressure signature from the Mach 2 wind-tunnel model: M = 2.0, h = 6 in, and

C L = 0.068; nacelles on.

The model's nacelles were thought to be a possible source of the shock, so they were removed and

another pressure signature, figure 3, was measured.

<1 0.000

ii!iiiii!i!

-.030

0.0

Figure 3. Pressure signature from the Mach

CL = 0.068, nacelles off.

!iiiiiii!iiiiiii

B.O 16.0

%,;n

2 wind-tunnel model: M = 2.0, h = 6 in, and

Comparisons of figures 2 and 3 showed that the unexpected strong shock observed in figures 2

was, as suspected, caused by the engine nacelles. However, the shock should not have been

present since the Mach 2 and Mach 3 models had been designed with the latest methods,

references 16 and 17, which should have accounted for these nacelle disturbances.

A methods analysis and re-evaluation revealed that the original methodology was based on

assumptions not completely fulfilled on real wind-tunnel models or real aircraft. These idealized

assumptions concerned the fineness ratio and inlet-lip angles of the engine nacelles: nacelle

diameter/length ratios of 1/10 or more, and inlet-lip angles of 1.0 degree or less. Nacelle diameter/

length ratios were actually closer to 1/8 and inlet-lip angles were in the 3.0 to 4.0 degree range.

The aerodynamic volume of the nacelle (integral of the nacelle cross-section area minus the inlet

area along the length) was usually a small contribution to the total equivalent area of the aircraft.

However, a lip angle of 3.0 to 4.0 degrees generated a strong shock and an abrupt "jump" in front

of the nacelle F-function because the inlet-lip radius is not zero as it is on the nose of a slender

3
bodyof revolution.Similarly,theareagradientat thestartof thenacelle-winginterference-lift

distributionhada linearcomponentwhichintroduceda similar"jump" in front of thenacellewinginterference-liftF-function.The analysisandits resultswerepresentedin reference19

whereit wasshownthat:(1) sharp-lip-anglenacelleswouldgenerateundesirableflow-field

shocksif thenacelleswerepositionedby low-dragratherthanby low-boomcriteria;and(2) the

nacellesgeneratedthesestrongshocksunderchoked-flowor unchoked-flowconditions.Soa

modificationwasmadeto thedesignandanalysismethod:equivalentareaswouldbesummedand

usedto calculateacombined-areaF-functiononly if eachareacomponentstartedwith zeroarea

andhadsmoothandcontinuousderivativesat theirorigins;F-functionsfor componentslike

enginenacelleswouldbecalculatedseparately,andthencombinedwith F-functionfrom the

componentswith smoothandcontinuousareadistributionsto obtainanaircraftF-function.

APPLICATION OF MODIFIED METHODOLOGY

The modified method for predicting sonic-boom overpressures and designing low-boom

conceptual aircraft seemed theoretically and empirically correct, but wind-tunnel tests were

necessary for validation. To achieve this objective, new aircraft concepts were designed,

references 20 to 22, and 1:300 scale wind-tunnel models were built. The concepts were designed

with the high-aerodynamic efficiency approach and methods of reference 23. They also

incorporated the results from the most recent low-boom technology studies. The desired goal was

to determine whether mission requirements and low-boom constraints could be met with aircraft

concepts whose gross takeoff weights were similar to those without sonic-boom reduction

features that were designed by United States aircraft companies. Mission characteristics were

determined by weights and performance codes similar to the code in reference 24. Sonic-boom

characteristics were predicted with Whitham theory during the design process or with Computer

Fluid Dynamics (CFD) methods for comparison with data from the wind-tunnel tests.

Two of the three conceptual aircraft, references 20 and 21, had engine nacelles in the usual

under-the-wing location where they generated both nacelle-volume and nacelle-wing

interference-lift disturbances. The third concept, the HSCT-10B, reference 22, had the engine

nacelles mounted on the aft-fuselage where they generated nacelle-volume, but not nacelle-wing

interference-lift disturbances, in the flow field. Large flow-field disturbances from the engine

nacelles would be readily observed in the measured pressure signatures from the HSCT-10B

model if the design methodology was inadequate or was improperly applied when the nacelles

were integrated with the wing-fuselage for reduced sonic boom. Such data would be helpful when

the complete body of wind-tunnel test data was interpreted.

The HSCT-10B concept was designed after an analysis of wind-tunnel data from the

previously-mentioned Mach 2 low-boom theory-validation concept, figure 1. Similar in length,

but with 6.25 percent less span, it was designed to have low sonic-boom characteristics at a Mach

number of 1.8 rather than 2. There were other performance-based design differences as well. The

HSCT-10B was designed to cruise at Mach 2.4 over water and Mach 1.8 over land. The Mach 2

theory-validation concept had the engine nacelles under the wings, but the HSCT-10B had its

engine nacelles mounted behind the center-section trailing edge on the aft fuselage. The Mach 2

concept had slightly-notched wing with an area of 15,000 ft 2, but the HSCT-10B had a

moderately-notched wing with an area of about 10,500 ft 2. A three-view sketch of the HSCT-10B

concept is shown in figure 4.

4
/

Figure 4. Three view of the aft-fuselage engine nacelle concept, the HSCT-10B.

The HSCT-10B wind-tunnel model had the sting integral with the fuselage. Strain gages were

recessed in the model sting to form a balance-sting support so that lift and pitching moment could

be monitored during the measurement of flow-field overpressures. This arrangement permitted a

smooth transition between the model and the sting support, but it meant that the area aft of the

fuselage altered the shape of near-field pressure signatures.

WIND-TUNNEL TEST DATA AND ANALYSIS

The HSCT-10B wind-tunnel model produced pressure signature data that were both

encouraging and disappointing. With the model at half the cruise lift coefficient, the top of the

measured pressure signature was markedly fiat and level. This was the desired shape that was

expected from the modified design, analysis, and nacelle-integration methods employed during

the design process. As the lift on the wind-tunnel model was increased, a slight "ramp" started to

form across the top of the positive-pressure section of the signature. As cruise lift was neared, the

aft end of the "ramp" erupted into a pressure disturbance which quickly grew to shock levels as

full cruise lift was reached. The appearance and strength of this shock was not expected.

However, these pressure signatures, with both large and small engine nacelles, somewhat

resembled the shape and features of the pressure signatures from the Mach 2.7 low-boom windtunnel models whose data was reported in reference 18. In the following sections, samples of

these pressure signatures from the HSCT-10B model will be presented. They were measured

through a range of C L from 0.5 CL,CRUISE to full CL,CRUISE. The practice of using varying lift

had been used before in previous wind-tunnel pressure-measurement tests with references 18, 25,

and 26 being typical though not an exhaustive set of examples.

HSCT-10B Concept Model The two measured pressure signatures in figure 5 and 6 are from

wind-tunnel tests, reference 27, of models derived from the HSCT-10B concept shown in figure 4.

They were measured at a Mach number of 1.8 and at a distance of 24 inches with the model at a

C L = 0.0511, which was a CL/CL,cRUISE ratio Of about 0.5.
.02

Ap

P

0.0

-.02 t I t I _ t I t I i

0 4 8 12 16 20

x, inches

Figure 5. Pressure signature from the HSCT-10B at M - 1.8, C L - 0.0511, and h - 24 inches;

small nacelles.

.O2

Ap 0.0

p

-.02 I I 1 ! I I ,,.[ I I

4 8 12 16

x, inches

I

2O

Figure 6. HSCT- 10B pressure signature: M = 1.8, C L = 0.0511, and h = 24 inches; large nacelles.

Both positive-pressure sections of the pressure signatures are virtually identical in shape, length,

and magnitude although the large nacelles were 19-percent longer and 39-percent wider than the

small nacelles. However, the C L was only half of CL,CRUISE. At this lift condition, the top of the

pressure signatures is virtually flat without any indication of wing-fuselage junction, wing

leading-edge crank, or nacelle shocks.

When the C L was increased to 0.1022, the cruise lift coefficient, the pressure signatures in

figure 7 and 8 were recorded.

.O2

Ap

0.0

P

-.02

I I [ I I I 1 I I 1

0 4 8 I2 16 20

z, inches

Figure 7. HSCT- 10B pressure signature: M = 1.8, C L = 0.1022, and h = 24 inches; small nacelles.
.O2

Ap

-- o.o

p

-.02

I I I---.,.,__1_ i I I t I j

0 4 8 t2 16 20

x, inches

Figure 8. HSCT-10B pressure signature: M = 1.8, CL = 0.1022, and h = 24 inches; large nacelles.

Again, both positive-pressure sections of the pressure signatures were virtually identical in

shape near the front with an additional strong shock (due entirely to wing lift) positioned before

the expansion to the tail shock. These changes in signature shape, from figure 6 to figure 8, over a

range of lift coefficients, CL/CL,cRUIS E of 0.50 tO 1.0, are shown in figure 9.

Lift/Design lift

0.500 ........

.O2 :'_ 0.625

/.A, 0.750 -- -- __

T o.o -- _'i: __:"

_, Z ¢

-.02 I , I ] I 1 I I [ •

0 4 8 i2 16 20

x, inches

Figure 9. HSCT-10B pressure signatures: CL/CL,cRUIS E ratios of 0.50, 0.625, 0.75, and 1.0;

M = 1.8, h = 24 inches, large nacelles.

Since the wing leading edge contour on the HSCT-10B planforrn was purposely made smooth

and continuous, the growth and strength of the lift-induced shock was a source of concern. This

rapid growth in flow-field disturbance was the result of a relatively small increase in angle of

attack. The positive-pressure segment on the signature in figure 5, CL/CL,CRUISE = 0.5, suggested

that, at the low lifting pressures across the wing upper and lower surfaces, the equivalent areas

and summation of F-functions represented the model's low volume and lift disturbances

reasonably well. With increasing lift, an ever larger pressure perturbation due to lift formed until,

at the designed-cruise lift coefficient, a strong shock, rather than a flat-topped pressure trace,

signaled the end of the pressure signature;s mid-section ramp and the beginning of the expansion

leading to the tail shock. This lift-induced disturbance grew into a strong shock over a range of
about1.5degreesin angleof attackwhich correspondedto anincreasein equivalentarea-due-tolift from 400 ft 2 to 800ft 2.

The growth and behavior of this lift-induced disturbance suggested that it was due to near-field

effects. If this hypothesis was correct, the observed extra shock was an imbedded shock that

would decrease in strength rapidly before it could move forward to coalesce with the nose shock.

The "ramp" section of the signature will lose its "roller-coaster" appearance and would approach

the shape of the desired ground signature. However, further wind-tunnel tests at larger separation

distances would be required to check this hypothesis.

The pressure signatures in figure 5 and 6 showed no large shocks from either the small or the

large engine nacelles, while the pressure signatures in figures 7 and 8 showed almost identical

shocks when the lift on the model was increased through the same range. These results were the

basis of a conclusion that the engine nacelles had been successfully integrated into the wingfuselage of the HSCT-10B concept and that nacelle-integration methodology could successfully

account for nacelle volume disturbances. Since only the small engine nacelles were used in the

design of the HSCT-10B, the carry-over to the large engine nacelles must be considered

fortuitous.

There was no time available to begin a new design that would test the complete nacelleintegration methodology. A "boom-softening" study was undertaken to reduce the sonic boom of

a configuration without employing the configuration tailoring described in reference 11. This

study, then, became an opportunity for employing and theoretically evaluating the full capabilities

of the Langley nacelle-integration methodology.

METHOD APPLICATION TO A "BOOM-SOFTENED" CONCEPT

The conceptual aircraft shown in figure 10 was designed as part of the "boom-softening" study

reported in reference 28.

Figure 10. Three view of the "boom-softened" arrow-wing concept, the HSCT-11C.

It differed from the baseline McDonnell-Douglas concept mainly in two ways: (1) its planform

8
had an extended and curved leading edge to "soften" the sonic boom without drastically

increasing the wing area and the gross takeoff weight; and (2) the configuration was given a

canard, rather than a horizontal tail, for takeoff and landing operations.

The design Mach number for the HSCT-11C was 2.4, the c_]se Mach number specified of all

the "softened-boom" study concepts. Figure 11 shows the complete F-function of the HSCT-11C.

F . .(}I

k(x.) / ,/ ". [ Nacelles off ....

0" , I I -. - I _ !

0 10(I 200

x e, ft

1 .1 •

300 400

Figure 11. F-functions of the HSCT-11C concept at M = 2.4 and W = 684,400 lb.

The location of the engine-nacelle disturbances, both volume and interference lift, are readily

apparent on these summed F-functions. In figure 12, the predicted ground pressure signatures,

nacelles on and nacelles off, are presented and compared for the HSCT-11C at cruise Mach

number and 56,850 ft.

3

6p, psJ

Nacelle_ on

I --"_- _ Nacelles orr

o ___

-3 I I I I I I I I I

-3(X) 0 300 (¢30

x - 13h,t'r

Figure 12. Predicted ground signatures from the HSCT-11C at M = 2.4, W = 684,400 Ib, and

h = 56,850 ft.

There is a 0.3 psf increment between the nacelles-on and nacelles-off condition which is

noticeable even on this non-low-boom-tailored configuration. If the engine nacelles could be

moved further aft, their disturbances would be superimposed on the F-function's expansion region

and considerably reduced in its effect on the nose-shock strength. The nacelles could be moved

further aft by extending the central wing section; an idea modeled after that feature on the Russian

Sukhoi S-51 conceptual aircraft, reference 29, which is shown in figure 13.
Figure 13.Three-viewof the SukhoiS-51conceptualaircraft.

This modificationto thewing planformof the HSCT-11C was employed as a "softened-boom"

feature on the HSCT-11E. With it, the wing-mounted engine nacelles could be moved aft so their

lip-shock disturbances should not add to the strength of the nose shock. The canard on the

conceptual HSCT- 11E, as on the HSCT- 11C, was used only for takeoff and landing. A three view

of the HSCT-11E concept is presented in figure 14.

Figure 14. Three view of the HSCT-11E concept (the modified HSCT-11C).

With the engine nacelles shifted to this aft location, the combined nacelle volume and interference

lift disturbances are also moved aft on the concept's F-function, as seen in figure 15.

10
.02 -

°,

F (x) ", Nacelles on

o.0

-.01

-.02

-.o3_

100 2O0 I J

3OO

xe, ft 400

Figure 15. F-functions of the HSCT-11E concept at M - 2.4 and W = 696,000 lbl

Predicted pressure signatures from the HSCT-11E concept, with and without nacelles, were

inobtained figure from 16. the F-functions in figure 15. These pressure signatures are presented and compared

3 h

Ap, p.q _ Nacelles on ---------

r 1-

-3, I _ _ • I I I I

-30 0 I

300

x -_h4"t 600

Figure 16.hPredicted = 56,850 ground ft. signatures from the HSCT-11E at M = 2.4, W = 684,400 lb, and

Although shifted aft over twenty feet, the nacelles still add a small incremental engine-nacelle

volume and interference lift disturbance to the strength of the nose shock. Moving the engine

nacelles further aft would have helped, but this would also have imposed an larger weight load on

the trailing edge of the wing's extended center section. However, the shape of the F-function in

figure 15 indicated that the wing planforrn shape and lifting length were the major contributors to

the nose-shock strength problem. If equivalent-area tailoring were applied to the concept's

volume and lift distribution by further alterations to the wing planform, the nose shock could be

reduced and the present location of the engine nacelles would not aggravate the reduced sonic

boom obtained by the use of low-boom methodology. Meanwhile, the extra inboard chord length

increased the available voIume for fuel and landing gear stowage If the wing thickness/chord ratio

was kept constant, there was additional spar depth for support of the engine nacelles. However,

there was an aerodynamic-performance decrement to be accepted since wing area was added

11
while thespanremainedconstant.Moreover,theadditionalwing areaandvolumewouldmeanan

increasein theconfiguration'sgrosstakeoffweight.

CONCLUDING REMARKS

An analytical and empirical technique for integrating the engine nacelles to a wing-fuselagefin configuration has been described. The technique was based on a more-complete application of

Whitham sonic-boom theory and sonic'boom minimization of Seebass and George, and was

developed to remedy the shortcomings of the previous method for analysis of sonic-boom

characteristics which was developed during the 1960's and 70's. Previous methods assumed that

engine nacelles slender with smooth and continuous area distributions, but the new technique

made no such assumption and permitted the application of low-boom methods along with high

aerodynamic efficiency methods to the analysis and configuring of a supersonic-cruise conceptual

aircraft during all stages of preliminary design.

Two HSCT concepts were used as examples to illustrate the application of this engine-nacelle

integration method: (1) a concept with engine nacelles mounted behind the wing on the aftfuselage, the HSCT-10B; and (2) a concept with engine nacelles mounted under wing centersection extensions, the HSCT-11E. In both cases, the key to reducing the flow-field disturbances

from the engine nacelles was to move them behind the wing where they would be in a "Machsliced" or area-ruled expansion region, where the wing lift was reaching its maximum value.

Mounting the nacelles behind the wing on the aft fuselage removed the interference-lift

disturbances and helped shift the center of gravity rearward, but it would place the nacelle inlets

in the upper-surface flow fields of the wing and the fuselage. It would also impose a large weight

and thrust loading on the tail section of the fuselage. Extending the wing center section so that the

engine nacelles could be mounted further aft under the wing moved both the nacelle volume and

the interference lift disturbances closer to the desired expansion region on the conceptual

aircraft's F-function. Additional benefits were extra wing volume available for fuel, landing gear

stowage, and nacelle support structure. However, these advantages were obtained at the expense

of extra wing area that reduced overall aerodynamic efficiency, and wing volume that would

probably increase the gross takeoff weight.

Both the Mach 2.0 theory-validation concept and the HSCT-10B were, theoretically, lowboom conceptual aircraft. However, if the "boom-softened" HSCT-11E would have been given

low-boom area-rule tailoring, it could have become an aircraft concept with engines in the underthe-wing location that possessed reduced- or low-boom characteristics. Nevertheless, both

concepts, the HSCT-10B and the HSCT-11E, demonstrated that low-boom technology and

nacelle-integration methodology along with high aerodynamic performance efficiency, innovative

structural efficiency, and advanced-composite materials could be included in the process of

preliminary conceptual-aircraft design. The HSCT-10B and the HSCT-11E concepts are most

likely not the only possible configurations that could achieve low-boom or "softened sonic-boom"

status. Since the canard on the HSCT-11E concept was intended only for takeoff and landing, both

concepts were flying wing configurations during the supersonic-cruise segment of the mission. It

is likely that there are canard-wing or wing-horizontal tail configurations, with the canard and the

horizontal tail carrying positive lifting loads, that might be achieve reduced- or low-boom flight

during supersonic cruise. In either case, the keys to achieving a low-boom conceptual aircraft

design would include the successful integration of engine nacelles on a wing-fuselage

12
configurationthat alsohasthepotentialfor high aerodynamicefficiencyat cruiseandlow speeds, andlow structuralweight.

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Publication 10133, 1993.

22. Mack, Robert J.: Low-Boom Aircraft Concept With Aft-Fuselage-Mounted Engine Nacelles.

High-Speed Research: Sonic Boom, Volume ]1, NASA Conference Publication 10133, 1993.

23. Baals, Donald D.; Robins, A. Warner; and Harris, Roy V. Jr.: Aerodynamic Design Integration

Of Supersonic Aircraft. No. 68-1018, 5 th AIAA Annual Meeting And Technical Display,

Philadelphia, Pennsylvania, October 21-24 1968.

24. McCullers, L. A.: Aircraft Configuration Optimization Including Optimized Flight Profiles.

Recent Experiences In Multidisciplinary Analysis And Optimization.

jarosiaw SObieski, compiler, NASA CP-2327, Part 1, 1984.

25. Carlson, Harry W.; McLean, E Edward; and Shrout, Barrett L.: A Wind-Tunnel Study Of

Sonic-Boom Characteristics For Basic And Modified Models Of A Supersonic Transport

Configuration. NASA TM X-1236, May 1966.

14
26.Hunton,Lynn W.; Hicks,RaymondM.; andMendoza,JoeIE: Some Effects Of Wing

Planform On Sonic Boom. NASA TN D-7-160, January 1973.

27. Mack, Robert J.: Wind-Tunnel Overpressure Signatures From A Low-Boom HSCT Concept

With Aft-Fuselage-Mounted Engines. High-Speed Research: 1994 Sonic Boom Workshop,

Configuration Design, Analysis, and Testing. NASA/CP-1999-209699, 1999.

28. Mack, Robert J.: Boom-Softening And Nacelle Integration On An Arrow-Wing High-Speed

Transport Concept. 1995 NASA High-Speed Research Program Sonic Boom Workshop,

Volume II - Configuration Design, Analysis, and Testing, NASA/CP-1999-209520, 1999.

29. Lambert, Mark, ed.: Jane's All The World Aircraft, 1994-95, 85 tlaEdition, Sentinel House,

Surrey UK. (Sukhoi S-51)

15
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December 1999 Technical Memorandum

4. TITLE AND SUBTITLE S. FUNDING NUMBERS

Nacelle Integration To Reduce the Sonic Boom of Aircraft Designed To

Cruise at Supersonic Speeds

6. AUTHOR(S)

Robert J. Mack

7. PERFORMING ORGANIZATION NAMEIS) AND ADDRESS(ES)

NASA Langley Research Center

Hampton, VA 23681-2199

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES

WU 537-07-21

W'BS 4.3.3

8. PERFORMING ORGANIZATION

REPORT NUMBER

L-17742

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/I_-1999-209534

12a. DISTRi6UTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE

Unclassified--Unlimited

Subject Category 05 Distribution: Nonstandard

Availability: NASA CASI (301) 621-0390

13. ABSTRACT (Maximum 200 words)

An empirical method for integrating the engine nacelles on a wing-fuselage-fin(s) configuration has been

described. This method is based on Thitham theory and Seebass and George sonic-boom minimization theory. With

it, both reduced sonic-boom as well as high aerodynamic efficiency methods can be applied to the conceptual

design of a supersonic-cruise aircraft. Two high-speed civil transport concepts were used as examples to illustrate

the application of this engine-nacelle integration methodology: (1) a concept with engine nacelles mounted on the

aft-fuselage, the HSCT-10B; and (2) a concept with engine nacelles mounted under an extended-wing center

section, the HSCT-I1E. In both cases, the key to a significant reduction in the sonic-boom contribution from the

engine nacelles was to use the F-function shape of the concept as a guide to move the nacelles further aft on the

configuration.

14.SUBJECTTERMS

Sonic boom; High-Speed Civil Transport; Supersonic aircraft

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OF REPORT

Unclassified

NSN 7540-01-280-5500

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Unclassified

19. SECURITY CLASSIFICATION

OF ABSTRACT

Unclassified

15. NUMBER OF PAGES

20

16. PRICE CODE

A03

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UL

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