NASATechnical Memorandum 4749 NASA-TM-4749 19960038270

_2

_ Wind-Tunnel Development of an

SR-71 Aerospike Rocket Flight

Test Configuration

Timothy R. Moes, Brent R. Cobleigh,Timothy R. Conners,

Timothy H. Cox, Stephen C. Smith, and Norm Shirakata

FORREFERE_

NOTTOBETAKENFROMTHISROOM

I.,.,

J

, June 1996 _.,/_!_._ :

I

[ i _" -,

National Aeronautics and

Space Administration

IIIIIIIIIIIIIIIII

3 1176 01426 1979

NASA Technical Memorandum 4749

_,, Wind-Tunnel Development of an

SR-71 Aerospike Rocket Flight

"rest Configuration

Timothy R. Moes, Brent R. Cobleigh,

Timothy R.Conners, and Timothy H. Cox

NASA Dryden Flight Research Center

Edwards, California

Stephen C. Smith

NASA Ames Research Center

Moffett Field, California

Norm Shirakata

Lockheed-Martin Skunk Works

Palmdale, Cafifornia

1996

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

....

WIND-TUNNEL DEVELOPMENT OF AN SR-71 AEROSPIKE

ROCKET FLIGHT TEST CONFIGURATION

Timothy R. Moes,* Brent R. Cobleigh,*

Timothy R. Conners,* and Timothy H. Cox t

NASA Dryden Flight Research Center

Edwards, California

Stephen C. Smith,*

NASA Ames Research Center

Moffett Field, California

" Norm Shirakata*

Lockheed-Martin Skunk Works

Palmdale, California

Abstract Nomenclature

A flight experiment has been proposed to investigate B.L. butt line, in.

the performance of an aerospike rocket motor installed CD drag coefficient in a lifting body configuration. An SR-71 airplane would

be used to carry the aerospike configuration to the CL lift coefficient

desired flight test conditions. Wind-tunnel tests were

completed on a 4-percent scale SR-71 airplane with the Cl rolling moment coefficient (stability axis)

aerospike pod mounted in various locations on the upper

fuselage. Testing was accomplished using sting and Cm pitching moment coefficient

blade mounts from Mach 0.6 to Mach 3.2. Initial test zero lift pitching moment coefficient

objectives included assessing transonic drag and Cm°

supersonic lateral-directional stability and control. Cn yawing moment coefficient (stability axis)

During wind-tunnel these tests, flight simulations were run with Cp surface pressure coefficient data to assess the acceptability of the

configurations. Early testing demonstrated that the Cy side forcecoefficient

initial configuration with the aerospike pod near the

SR-71 center of gravity was unsuitable because of large c.g. center of gravity

nosedown pitching moments at transonic speeds. The DFRC NASA Dryden Flight Research Center,

excessive trim drag resulting from accommodating this Edwards, California

pitching moment far exceeded the excess thrust

capability of the airplane. Wind-tunnel testing continued ES. fuselage station, in.

in an attempt to find a configuration suitable for flight KEAS equivalent airspeed, kn

test. Multiple configurations were tested. Results

indicate that an aft-mounted model configuration LASRE Linear AerospikeSR-71 Experiment

possessed acceptable performance, stability, and control M Mach number

characteristics.

m.a.c, mean aerodynamic chord

NTS National Technical Systems

*Aerospace Engineer. AIAA member. SSTO single stage to orbit

tAerospace Engineer.

:_Retired Aerospace Engineer. W.L. waterline

Copyright © 1996 by the American Institute of Aeronautics and

Astronautics, Inc. No copyright is asserted in the United States under O_ wing reference plane angle of attack, deg

Title 17, U.S. C_xle. The U.S. Government has a royalty-free license to

exercise all rights under the copyright claimed herein for Governmental [_ angle of sideslip, deg

purposes. All other rights are reserved by the copyright owner. A increment from baseline SR-71 aircraft

1

American Institute of Aeronautics and Astronautics
8e elevon deflection, deg (fig. 1). Altitude compensation is a result of the plume

8r rudder deflection, deg being able to expand to the slipstream static pressure.

The slipstream pressure, however, is necessarily reduced

Aerodynamic Model Reference Quantities because of flow expansion around the rocket nozzle

cowl. This decrease in pressure causes the plume to

bref reference span, 56.7fl overexpand slightly; hence, an expected loss of

Cref reference chord = m.a.c., 37.7 ft performance occurs. This phenomenon has been labeled the "slipstream effect" and is dependent on the

Sr_f reference area,1605 ft2 configuration, Mach number, chamber pressure ratio,

Xr_f longitudinal moment reference, F.S. 900 and Reynolds number. .

(25 percent m.a.c.) In an effort to obtain data for the X-33 concept,

Yref lateral moment reference, B.L. 0 Lockheed-Martin and Rocketdyne, Canoga Park,

California (builder of the aerospike rocket), initiated a

Zr_ vertical moment reference, W.L. 100 program with NASA to flight test the aerospike rocket

Introduction engine using an SR-71 airplane as the carrier vehicle in

a project known as the Linear Aerospike SR-71

Recent emphasis on single-stage-to-orbit (SSTO) Experiment (LASRE). The aerospike rocket and its

vehicle technology has led the National Aeronautics and associated components, referred to as the aerosl)ike pod.

Space Administration (NASA) to request proposals for will be mounted on the SR-71 upper fuselage surface

an SSTO Technology Demonstrator Flight Vehicle centerline. Figure 2 shows the pod installation on the

which wind-tunnel model. Note that the vertical tails of the has become known as the X-33. A contractor

team, led by Lockheed-Martin Corporation, Marrietta, SR-71 airplane were removed for this photograph. The

Georgia, developed an X-33 concept which incorporates aerospike pod components are referred to as the canoe,

a linear aerospike rocket engine in a lifting body vehicle kayak, reflection plate, and a half span model of a lifting

configuration. The aerospike rocket was first developed body configuration. The canoe is installed on the SR-71

and ground tested in the 1960's. 1,2 The principle fuselage and contains the gaseous hydrogen fuel and

advantage of the aerospike rocket is the inherent altitude liquid water needed for cooling. The kayak, located

compensation provided by the nozzle. This beneath the reflection plate and on top of the canoe, is

compensation results in increased specific impulse used to incline the model at a 2° nosedown incidence

performance during the low-altitude portion of a flight angle. This angle aligns the lower part of the model with

when compared to a conventional bell nozzle rocket the local flow over the top of the SR-71 airplane. The

Aeroske'

nozzle-__ jj

Specific J JJJ_

impulse //

j Bell

.. 4" " nozzle

_ _. - _ _" Bellnozzle Aerospikenozzle

Sealevel Highaltitude

Altitude

960227

Figure 1. Qualitative comparison of aerospike and bell nozzle specific impulse.

2

Amcricm_ Institute of Aeronautics arid Astronautics
L

Figure 2. Aerospike pod wind-tunnel installation.

reflection plate is mounted on top of the kayak to determine if the configuration could meet the test

promote uniform flow in the region of the model. Liquid objectives and to develop flight techniques for the

oxygen and ignitor materials required to operate the transonic acceleration that would minimize fuel usage.

engine are located in the model. The model is mounted

vertically so that sideslip of the SR-71 airplane imparts Stability and control issues investigated in the wind

angle of attack on the model. Unlike an actual X-33 tunnel were primarily lateral--directional at takeoff and

configuration, the model has no camber so as not to landing speeds and at the maximum Mach number. At

produce side force when the SR-71 airplane is at low speeds, the cross-wind landing criterion with one

0° sideslip. The desired Mach number and altitude engine not functioning needed to be satisfied with the

ranges for the test extend to Mach 3.2 and an altitude of aerospike pod configuration. At supersonic speeds,

84,000 ft which are near the limits of the basic SR-71 the SR-71 lateral-directional stability minimum occurs

flight envelope, at Mach 3.2. Ensuring that the aerospike pod

configuration does not significantly reduce the stability

A series of wind-tunnel tests was conducted to at Mach 3.2 was necessary. The incremental stability

support this flight project. The objectives were to define and control derivatives were added to the flight

a LASRE configuration that could successfully be flown simulations to assess handling qualities and envelope

to the desired test points and to identify any aircraft restrictions caused by control surface authority a

envelope limitations. Principle areas of concern were limitations.

transonic drag, supersonic stability, and control.

Aerodynamic increments obtained in the wind tunnel This report presents selected results from three windrepresent the aerodynamic changes caused by the tunnel entries and flight simulation results based on

addition of the LASRE pod to the baseline SR-71 wind-tunnel data. Potential flow analyses are also

airplane, presented to help interpret the wind-tunnel results. The

discussion of the results includes the unexpected

The excess thrust available for the SR-71 airplane is nosedown pitching moment encountered with the first

minimal at transonic speeds. As a result, minimizing the configuration tested and the solution to the pitching

transonic drag caused by the aerospike pod was moment problem by altering the location of aerospike

necessary. The incremental drag obtained in wind- model; drag, stability, and control of various

tunnel testing was incorporated into simulations of configurations; envelope limitations for the LASRE

complete flight profiles. These simulations were done to configuration; and maximum Mach number

3

American Institute of Acronautics and Astronautics
60 -- Wind

tunnel

50 / _ --O-- 7ft

"_ --E}-- 4ft

Flight I

40 . I

test I

Reynolds

number, 30 -

million

20

10

o=P- I-0 I

.5 1.0 1.5 2.0 2.5 3.0 3.5

Machnumber

960229

Figure 3. Comparison between wind tunnel and flight Reynolds numbers (based on aerospike model length).

performance expected in flight. Use of trade names or reflected shocks and to set subsonic and transonic Mach

names of manufacturers in this report does not numbers. The transonic cart was removed for Mach 1.6

constitute an official endorsement of such products or and above. Supersonic Mach number was controlled by

manufacturers, either expressed or implied, by NASA contouring the upstream nozzle. At 8° angle of attack, ct,

wind-tunnel blockage was approximately 4.7 percent.

Wind-Tunnel Facilities

The model was supported with a sting mount entering

Two wind-tunnel facilities were used to test the the aft end of the fuselage. The force and moment

LASRE configuration: the National Technical Systems balance was aligned with the fuselage reference plane of

(NTS) 4- × 4-ft blowdown wind tunnel in Saugus, the model. Angle of attack was referenced to the wing

CaliforniJ and the MicroCraft 7- × 7-ft blowdown reference plane which has an incidence angle of-1.2 °

wind tunnel in El Segundo, California.** Minimum with respect to the fuselage reference plane. A standard

operating Reynolds numbers were adequate for these six-component balance was used for Mach 2.0 and

tests; therefore, these Reynolds numbers were used to below.

conserve tank pressure and to maximize the time

available for testing. Figure 3 shows a comparison of A five-component instrumented sting was specially

wind-tunnel and flight Reynolds numbers data. The fabricated and used from Mach 1.6 to Mach 3.2. The

model installation and test procedures for each facility instrumented sting was necessary because the standard

are described next. balance was not expected to be able to withstand the

wind-tunnel start-up loads at the high Mach numbers.

NTS 4- × 4-fl Blowdown Wind Tunnel The five-component instrumented sting did not have an

axial force measurement; therefore, lift and drag could The model was tested from Mach 0.6 to Mach 3.2. A

not be computed.

transonic cart was inserted in the wind tunnel for testing

at Mach 1.4 and below. The 22-percent porous walls of The majority of the longitudinal aerodynamics data

the transonic cart were actively suctioned to minimize was obtained during angle-of-attack sweeps at 0°

sideslip. Some angle-of-attack sweeps were also done at

_Wind Tunnel User's Manual. Kelly JohnsonResearch and -5 ° sideslip. Lateral-directional data were obtained

DevelopmentCenteratRye CanyonLockheedAeronauticalSystents during angle-of-sideslip sweeps with an angle of attack

Company,FluidDynamicsLaboratory,Jtme1990.[NTS;RyeCanyon of 4° (approximate cruise flight value). Research,Development,andTest Center;25100RyeCanyonRoad;

Building202;Valencia,California91355,Telephone:(805)259-8184,

FAX:(805)257-35391 MicroCraft 7- x 7-ft Blowdown Tunnel

**Melmell,R.C.,MicroCraftTrisonicWindTunnelUser'sManual,

NA-90-1484,MicroCraft,Inc.,El Segundo, Califomia,Jan. 18,1994. The model was tested from Mach 0.6 to Mach 1.6. All

[MicroCraft,Inc.,TrisonicWindTunnel;400DuleyRd.;ElSegundo, testing was done in the transonic section which had

California90245,Telephone:(310)335-1585,FAX:(310)640-1056] 19.7-percent porous walls. Mach number was controlled

4

AmericanInstituteof AeronauticsandAstronautics
in the test section by adjusting the flexible nozzle, the to as the forward model and aft model (figs. 4 and 5).

downstream diffuser configuration, and a Mach number The majority of the pod components were fabricated of

control flap which operated at 40 Hz. At 8° angle aluminum. The lone exception is the reflection plate

of attack, wind-tunnel blockage was approximately which was fabricated of stainless steel. Two canoes were

1.5 percent, used. The first canoe was used with the forward model

configuration, and the slightly longer second canoe was

The model was supported by a blade mount attached used with the aft model configuration. Use of the kayak

to the fuselage underside. A six-component balance was was limited to the aft model configuration. Grit strips

installed in the model to obtain the aerodynamic loads, were used on the pod to create a turbulent boundary

Mach number was limited to Mach 1.6 because of wind- layer over the pod components at approximately

tunnel start-up loads on the balance. Angle-of-attack 5-percent chord. The grit grain was sized to 0.0034 in.

sweeps were used to obtain longitudinal data. Angle-of- for all tests.

.. attack sweeps at -5 ° sideslip were used to obtain

lateral-directional data because angle-of-sideslip Wind-Tunnel Testing

sweeps were not possible using the blade mount.

Table 1 summarizes the wind-tunnel test sequence.

Wind-Tunnel Models The first wind-tunnel entry occurred at the NTS 4- x 4-ft

tunnel. The initial LASRE configuration tested was the

An existing 4-percent scale model of the SR-71 forward model configuration (fig. 4). After testing

airplane was used for these tests. The main fuselage and numerous configurations, the final configuration tested

wings were constructed of cast bronze. Aluminum was was an initial aft model configuration. Because of the

used for the forebody and tails. Flow through inlets were limited amount of wind-tunnel time remaining, only a

used for the model. The aft fuselage and inboard elevons partial set of aerodynamic data was obtained for this

were modified to accommodate the sting mount at the configuration. Before the second wind-tunnel entry, the

4-ft tunnel. The true aft fuselage was used at the 7-ft aft model configuration was modified slightly by

tunnel test in which the model was blade mounted. The extending the length of the canoe and adding the section

vertical tails were manually positioned with brackets for known as the kayak to incline the reflection plate 2°

0 °, 5 °, and 10° rudder deflections. Separate inboard nosedown (fig. 5).

elevons were fabricated for-10 °, 0°, and 10° elevon

deflections. The second wind-tunnel entry occurred at the

MicroCraft 7- x 7-ft wind tunnel because it could

The aerospike pod consisted of a canoe, reflection accommodate a blade mount. The blade mount was

plate, model, and, in some cases, a kayak (fig. 2). The necessary because testing the configuration with the true

pod was tested in various locations on the SR-71 SR-71 aft fuselage was desired. Testing was primarily

airplane. The two primary configurations were referred on the aft model configuration from Mach 0.6 to

Table 1. Wind-tunnel test summary.

Mach

Entry Facility Mount Balance number Objectives

- Transonic drag

1 NTS Sting Six component 0.6-2.0 Stability

Rudder effectiveness

Five component 2.5-3.2 Stability Rudder effectiveness

Transonic drag

2 MicroCraft Blade Six component 0.6-1.6 Stability Rudder effectiveness

Elevon effectiveness

Stability

3 NTS Sting Five component 1.6-3.2 Rudder effectiveness

Elevon effectiveness

5

AmericaJl Institute of Acronautics and Astronautics
F.S.934.5 1

F.S.754.5 -1 / 57 8 in r-F.S.1226.5

F'S'736"5 -] / o K /- W.L.154.9 /'_--_

W.L.131-_ F .....

.900j ' Sting mount-_F.S:;29_:_

F S 1355--'

B.L.45-- __ _6.5

....

F.S.1295-

F.S.1355-

960230

Figure 4. LASRE forward model configuration.

Mach 1.6. Data were also obtained for the forward Aerodynamic Model

model configuration at Mach 1.I with the modified and

true SR-71 aft fuselages. For this entry, elevon Aerodynamic data for the baseline SR-71 and LASRE

effectiveness was also tested, configurations were obtained in these wind-tunnel tests.

Increments in the aerodynamic coefficients and 

The final entry occurred at the NTS 4- x 4-ft wind derivatives were calculated by subtracting the windtunnel. The instrumented sting was used for testing the tunnel baseline SR-71 data from the wind-tunnel data

aft model configuration from Mach 1.6 to Mach 3.2. The for the LASRE configuration. These increments were

standard balance was not used for Mach 2.0 and below then added to the baseline SR-71 rigid body

to save wind tunnel set up time. In addition, drag was aerodynamic model 3 to obtain the LASRE aerodynamic

not a required measurement for these tests. Elevon model. This model was then used in the simulations for

effectiveness was again tested with special elevons performance, stability and control, handling qualities

fabricated to accommodate the sting mount, analysis, mission planning, and crew training.

6

American Institute of Aeronautics and Astronautics
- F.S.1196.5

- F.S.1023.0 -F.S.1230.9

IFF.S-1033.0

F.S.1041.9

W.L.131• - ---iFF.S.749.6 ._J,...............FF.S.900 IIL...... _ _l....._ -,___ -r . 20

F S 102 '_ StingmountJ

• " F.S.1244.6-' | I

F'S'I2Fg._355 ._ J

960231

Figure 5. LASRE aft model configuration.

Flexibility corrections to the rigid body aerodynamics are a function of Mach number; whereas, others are a

of the baseline SR-71 airplane are documented in function of Mach number and angle of attack.

reference 3 and included in the simulation. No attempt

was made to identify the flexibility corrections caused CL + ACL(M, o_)

by the addition of the LASRE pod. The force and = CLBaselinc

moment equations used in the simulations are shown

below with the inclusion of the increments measured CD + ACo(M, o_)

during wind-tunnel testing. The form of these equations = CDBaseline

was determined as a result of the wind-tunnel testing.

The only increments included in the aerodynamic model Cm = + ACre(M, oO+ACm_e(M)Se

were those needed to adequately represent the LASRE CmBaseline

configuration aerodynamics. Some of these increments

7

Americanhistituteof AeronauticsmidAstronautics

r
Cy = CYBaseline + ACyfl(M, 0_)_ + ACY_r(M)6r instrumented sting was 1 percent of applied load. Good repeatability of these data within this accuracy was

demonstrated at the 7-fl tunnel. Repeatability was not as

C l -- C/Baseline + ACId(M, o_)_ + ACIsr(M)6r good at the 4-fl tunnel when comparisons were made of

runs from different days.

Cn = CnBaseline + ACn[_(M, o_)_ + ACn_r(M)_)r On the other hand, good repeatability was obtained in

the 4-ft tunnel for data runs that occurred during same

day testing. In terms of force and moment accuracies at

The reference area for the SR-71 airplane is 1605 ft2, this wind tunnel, the worst repeatabilities resulted in

the reference chord is 37.7 ft, and the reference span is ACm differences of 0.006, ACL differences of 0.03,

56.7 ft. The moment reference is located at fuselage and AC o differences of 0.0025. The force and moment station (ES.) 900, waterline (W.L.) 100, and butt line

(B.L.) 0. balance results presented here are more nearly accurate

than these quoted differences because wherever possible

•Wind-Tunnel Comparisons increments were obtained from same day testing.

A limited number of data comparisons between the

two wind-tunnels was made. Of particular interest were .o14 -- Wind

the transonic results because reflected shocks and tunnel tunnel

blockage in these small wind tunnels were a concern. .012 -- 7 it

Figure 6 shows comparisons for the aft model ,lit

configuration. The transonic data in the 7-ft tunnel were .OlO

taken with the final aft model configuration. The

transonic data in the 4-ft tunnel came from the first wind ACo .ooa

tunnel entry before the final configuration was

established. The configuration used in the 4-ft tunnel .o06

test did not include the kayak, and the reflection plate

was not inclined 2° . The reflection plate was raised .004

approximately 3/16 in. at three post locations. In 4°angle°f attack

addition, the length of the canoe was slightly shorter in .002 ] ] I ]

the 4-ft tunnel test. The aerodynamic effect of these .5 1.0 1.5 2.0 2.5

differences was assumed to be minor. Machnumber 960232

(a) Untrimmed drag increment.

The untrimmed drag data showed excellent agreement

between the two tunnels for supersonic Mach numbers

(fig. 6(a)). The 7-ft wind tunnel showed significantly .005 --

larger drag increments for Mach 0.9 and Mach 0.95 than

the 4-ft wind tunnel. The reason for this difference is

unknown. Because blockage was low for the 7-ft tunnel o --

and these results were conservative, the drag increments

from this 7-ft tunnel were used in the subsequent -.005 --

performance studies. Figure 6(b) shows the pitching ACm Wind

moment increments at 4 ° angle of attack for the two -.010 tunnel

wind tunnels. The shapes of these curves agree well --o-- 7ft "

as seen in the peak nosedown pitching moment at ----or- 4it

Mach 0.9 and the peak nose-up pitching moment at -.o15 --

Mach 1.05. 4° angle of attack

-.o2o I I I I

Wind-Tunnel Measurement Accuracies .5 1.0 1.5 2.0 2.5

Machnumber

960233

The expected accuracies of the internal sixcomponent balances used were 0.25 percent of applied (b) Pitching moment increment.

load. The expected accuracy of the five-component Figure 6. Comparisons of data from the 4- and 7-ft wind

tunnels for the LASRE aft model configuration.

8

AmericanInstituteof AcronauticsandAstronautics
The majority of the flow angle accuracies were well fuel use during the acceleration through the transonic

within 0.2 °. The 7-ft tunnel flow angle data were Mach numbers and, therefore, maximize the final Mach

expected to be within 0.1 o because there was not much number attainable for testing was of particular interest.

deflection of the blade mount. On the other hand, some

"slop" in the installation was attributed to an unknown Results and Discussion

error of _+0.05°. For the sting-mounted tests at the 4-ft

tunnel, additional uncertainty existed because of fairly This section presents representative wind-tunnel

large sting deflections, results and some comparisons with potential flow

results. The forward model configuration is discussed

_' Potential Flow Computer Tools first because it was the initial LASRE configuration.

This configuration was unacceptable because of an

A computational analysis of the LASRE configuration excessive amount of nosedown pitching moment at

. was conducted to validate the wind-tunnel results and to transonic speeds. Numerous configuration investigations

allow diagnosis of the aerodynamic phenomenon revealed that the aft model configuration was

associated with many of the LASRE configurations acceptable. A discussion of the aerodynamics associated

tested. The two codes used in this study are commonly with the aft model configuration is followed by a

known as A5024 and TranAir. 5,6 The A502 code, a description of the causes of the nosedown pitching

linear potential (PrandtI-Glauret) flow solver, was run moment. This section concludes with the ramifications

on a workstation and is useful for subsonic and of the aerospike pod on SR-71performance and control.

supersonic flows but not for transonic. TranAir is a

finite-element, full-potential flow solver and can, Forward ModelConfiguration

therefore, be run at transonic Mach numbers. TranAir Figure 4 shows the forward model configuration.

was run on a Cray C-90 (Cray Research, Inc., Before wind-tunnel testing, this configuration was the

Minneapolis, Minnesota) computer at the NASA Ames baseline for the LASRE project because of center of

Research Center, Moffett Field, California. The input gravity (e.g.) and flow quality considerations. The grid for A502 and TranAir is identical. aerodynamic model increments obtained for the forward

Flight Simulation model configuration are tabulated in appendix A.

Flight simulations using the LASRE aerodynamic Longitudinal aerodynamic increments are tabulated in tables A-l, A-2, and A-3. Trim angle of attack for the

database were integral to developing the final SR-71 airplane is roughly 4 ° for the majority of the

configuration. These simulations assessed performance, flight Mach numbers. Figure 7 shows the untrimmed stability and control, and handling qualities associated drag and pitching moment increment curves for 4 ° angle with the various configurations. Flight simulations were

conducted concurrently with wind-tunnel testing, and

the results were used to modify the wind-tunnel test

.014 --

plan. t_

Three simulations were used at DFRC during .012 7Xt:_

development of the LASRE configuration. These .010-

simulations are referred to as the batch sim,

per)brmance sim, and real-time sim. The batch sim was .008

run on a workstation for a prescribed flight trajectory. AeD

Outputs from the batch sim were used for stability and .006

control, handling qualities, operational limits, and .004

performance analyses. The trimmed drag outputs from

the prescribed trajectory were the primary input to the .002 4° angle of attack

performance sim. The performance sim was also run on ] ] ] ]

a workstation and was used to assess transonic and 0

.5 1.0 1.5 2.0 2.5

maximum Mach number performance of the LASRE Mach number

configuration. The real-time sim is a full-cockpit, _o_34

piloted simulation in which project pilots could develop (a) Untrimmed drag increment.

flight trajectories and assess handling qualities. Figure 7. Aerodynamic increments caused by the aeroDeveloping a flight trajectory that would minimize the spike pod with the model in the forward position.

9

American Institute of Aeronautics and Astronautics

........ -'--r .......
o -- Hinge moment analysis also showed that there was not

enough actuator control power for trim.

-.005 Lateral-directional stability of the forward model

configuration was investigated in the 4-ft tunnel. The

model was swept through +8° angle of sideslip with the

-.OLO angle of attack set to a value near the expected trim

ACre angle of 4°. Data were obtained at four Mach numbers

-.o15 (table A-4).

Figure 9 shows the yawing and roiling moment

-.020

4° angle of attack sideslip derivatives for the baseline SR-71 airplane and

-.026 [ [ [ the forward model LASRE configuration. The Cn is

cr 15

.s 1.o 1.s 2.0 significantly reduced for the LASRE confi,_uration and

Mach number

_0235 rapidly approached instability as the Mach number

(b) Pitching moment increment, increased. This confguration would have resulted in the

Figure 7. Concluded. LASRE being limited to a maximum of approximately

Mach 1.9 because of decreased directional stability. The

of attack. These data were obtained in the 4-ft wind rolling moment derivative, Clf_,

demonstrated an

unexpected decrease in rolling stability. It was thought tunnel. The peaks in the drag and pitching moment

occur at Mach 1.1. For 4 ° angle of attack, the maximum that this derivative would become more negative

untrimmed drag increment is 128 drag counts (fig. 7(a)), because of the large lifting body above the SR-71

and the maximum nosedown pitching moment centerline.

increment is 0.0208 (fig. 7(b)). This amount of

nosedown pitching moment is excessive, and the cause Aft Model Configuration

is discussed in the Nosedown Pitching Moment

subsection. Figure 5 shows the aft model configuration. The

complete set of aerodynamic model increments

Figure 8 shows the trim drag penalty associated with

the forward model configuration. The flight simulators

showed that the combined pod and trim drag increment

was greatly in excess of the power available for --o-BaselineSR-71 Cil3

acceleration through the transonic Mach numbers. ----o-BaselineSR-71Cnl 3

--0--- LASRE forward configuration ell _

--B-- LASRE forward configuration Cnl 3

.012 -- .003 --

.008 -- .__.m n_ __table

le of attack Stability .001 ll.--"'m _-_O--_ un_

derivative,

ACD .006 -- deg 0 --

.004 -- angle

-.001 e

.002o _q_ I I I -o02 I I I I.5 1.0 1.5 2.0 2.5 3.0

.5 1.0 1.5 2.0 2.5 Mach number 960237

Mach number

_0236 Figure 9. Yawing and rolling moment sideslip derivaFigure 8. Trim drag increment caused by the aerospike tives for the baseline SR-71 airplane and the LASRE

pod with the model in the forward position, forward model configuration.

10

American Institute of Aeronautics and Astronautics
measured for the aft model configuration is tabulated in untrimmed drag peak from Mach 1.1 to Mach 0.95

appendix B. During the first wind-tunnel test at the 4-ft (fig. 10(a)). For all supersonic Mach numbers, the

tunnel, an initial aft-mounted model configuration was untrimmed drag increment at 4° angle of attack was

tested which did not include the kayak or the 2° smaller for the aft model configuration. The large

nosedown incidence. In addition, the reflection plate improvement obtained in the pitching moment

was raised approximately 3/16 in. above the canoe, increment was more important than the drag. Above

Mach 0.9, the pitching moment increment was much

Figure 10 shows a drag and pitching moment closer to zero with the aft model (fig. 10(b)). A

comparison between the forward- and aft-mounted somewhat increased nosedown pitching moment was

_, configurations. Moving the model aft .changed the measured for Mach 0.6 and Mach 0.9.

As in the forward model configuration, the drag and

Model pitching moment increments for the aft model were

configuration incorporated into the batch simulator to come up with

--O--- Forward

.014 _ Aft trimmed drag estimates for a nominal LASRE flight and

c.g. profile. Figure 11 shows the trimmed drag estimates

.o12 _x_ for the forward and aft model configurations. The

increased drag increments at Mach 0.9 and Mach 0.95

.OlO initially were a concern; however, studies in the

performance simulator showed that the affect of these

.oo8 increased drag increments was not of major significance

ACD because of sufficient excess thrust of the SR-71 airplane

.006 in this Mach number range.

.004 Nosedown Pitching Moment

.o02 4° angle of attack The large nosedown pitching moment associated with

0 I ] I the forward model configuration was unexpected. The

.5 1.0 1.5 2.0 primary effect of the pod was expected to be a drag

Mach number _o23a component acting above the centerline which would

(a) Untrimmed drag increments, produce a nose-up pitching moment. A literature review

subsequently revealed that this nosedown phenomenon

Model

configuration

--O--- Forward -O-BaselineSR.71

.oos -- ----1:3-- Aft -El-Aft modelconfiguration

.050 -_-Forwardmodelconfiguration

o

.045

--lOGS .040

ACm -.010 .035 --

CD

-.015 .030

-.020,025 I 4°IangleofattackI .025 .020-- 40angle __%o°f attack _"-"_

.5 10 1.5 20 I I I I I I I

Mach number .015 96o2a9 .8 .9 1.0 1.1 1.2 1.3 1.4 1.5

(b) Pitching moment increments. Machnumber 960240

Figure 10. Comparison of aerodynamic increments Figure 11. Trimmed drag comparison between forward

caused by the aerospike pod with the model in the and aft model configurations.

forward and aft positions.

11

American Institute of Aeronautics and Astronautics

i •
was not without precedent. A search of the Lockheed- the nosedown pitching moment for angles of attack

Martin archives uncovered a wind-tunnel test in which a greater than 0°. For angles of attack greater than 3°, onetop-mounted payload caused a similarly large nosedown half of the nosedown pitching moment was caused by

pitching moment. 7 the canoe, and one-half was caused by the model.

Similar results were shown in the component buildup at

Because trim drag is a key performance issue in the Mach 1.6 (fig. 13). At Mach 1.6, the model added a

design, an effort was made to understand the causes of approximately equal to that of the canoe.

the nosedown pitching moment. A component buildup Cm°

of the configuration and an investigation into the aft Aft Fuselage Effects

fuselage effects were conducted in the wind tunnel.

Potential flow analyses were also done to help For the first wind-tunnel test, the aft fuselage of the

understand this phenomenon. SR-71 airplane was modified to accommodate the sting

installation. Uncertainty existed as to whether or not the

Component Buildup modification contributed to the large nosedown pitching

moment. The second wind-tunnel test provided an

The additive pitching moment effects of the opportunity to determine if there was an effect because

individual components of the forward model the model was blade-mounted and had the true aft

configuration were measured at Mach 1.1 and Mach 1.6 fuselage. The baseline SR-71 airplane and forward

in the 4-ft tunnel. The component buildup consisted of model configurations were tested with the actual and

the baseline SR-71 airplane and then individually modified SR-71 aft fuselages.

adding the canoe, reflection plate, and model. For

Mach 1.1, the addition of the canoe alone was Figure 14 shows the results for Mach 1.1.

responsible for the majority of the negative Cmoshift Figure 14(a) shows the pitching moment as a function of

(fig. 12). angle of attack, and figure 14(b) shows the pitching

moment increments. The aft fuselage modification did

The team speculated that for supersonic Mach not significantly affect the nosedown pitching moment

numbers, the shock on the bow of the canoe caused a increment at any angle of attack (fig. 14(b)).

high-pressure region forward of the moment reference

(ES. 900). In addition, expansion fans on the aft end of Potential Flow Analysis

the canoe caused a low-pressure region far aft of the

moment reference. Both of these effects would TranAir analyses were done at Mach 1.1 to assess the

contribute to a nosedown pitching moment. The influence of the canoe and the SR-71 aft fuselage on the

addition of the reflection plate had minimal affect on the pitching moment. In the component buildup phase of

pitching moment. The addition of the model increased the wind-tunnel test, results were obtained for the

.02 I_ -O-BaselineSR-71 .03 I_ -O-BaselineSR-71

-_-Canoe only I _ -_-Canoe only

.01 _ × Canoeandreflectionplane .02 × Canoeandreflectionplane

0 ition .01

-.01 0

Cm -.02 -- Cm -.01

-.03 -.02 --

-.04 -.03

-.05 -.04

-05 I I I I t t -.o5 I I I t I-=1

-2 0 2 4 6 8 10 -2 0 2 4 6 8 10

Angleof attack,deg 96o241 Angleof attack,deg 960242

Figure 12. Pitching moment coefficient for the forward Figure 13. Pitching moment coefficient for the forward

model configuration component buildup at Mach 1.1 in model configuration component buildup at Mach 1.6 in

the 4-ft wind tunnel, the 4-ft wind tunnel.

12

AmericanInstituteof AeronauticsaaldAstronautics
-O-BaselineSR-71(trueaftfuselage) --B-- Windtunnelwithmodifiedaftfuselage

-El-Forwardmodel(trueaftfuselage) [] TranAIrwithmodifiedaftfuselage

-<>-BaselineSR-71(modifiedaft fuselage) <> TranAIrwithtrueSR-71aft fuselage

× Forwardmodel(modifiedaft fuselage) 0 --

.04 -- _

.02 -.005

0 ACm-.010 _ -- _" _- -- 1,,

Cm

-.02 --

-.015 --

-.04 --

-.o6 I I I I -020 I I I I I

-2 0 2 4 6 8

- 4 - 2 0 2 4 6 8 Angleofattack,deg 960245

Angleof attack,deg 960243 Figure 15. Comparison of TranAir and wind-tunnel

results for the canoe-only configuration at Mach 1.1. (a) Pitching moments.

o TrueSR-71aftfuselage Figure 16 displays TranAir surface pressures for the

El ModifiedSR-71aftfuselage baseline SR-71 airplane and canoe-only configurations

o -- at 4° angle of attack. A high-pressure region exists

forward of the canoe, and a low-pressure region exists

aft of the canoe as compared to the baseline SR-71

-.005 -- airplane (figs. 16(a) and 16(b)).

[] These pressures can also be seen in the fuselage

ACm-.010 -- O surface pressure distribution for B.L. 19.4 (which is

gl 8 slightly outboard of the canoe outer wall at B.L. 15.0)

B (fig. 17). Recalling that the moment reference is at

-.015 -- [] F.S. 900 easily clarifies how the canoe alone can cause a

B ,., nosedown pitching moment. A shock is formed off the

I I I ._[ o _ _ [ bow of the canoe, pressurizing the SR-71 fuselage

[] forward of the canoe. In addition, an expansion fan is -4 -2 0 2 4 6 8

extended at the end of the canoe, lowering the pressure Angleof attack,deg 960244

on the fuselage.

(b) Pitching moment increments.

Figure 14. Comparison of pitching moment effects Unlike the supersonic data, wind-tunnel results

.. caused by the true SR-71 aft fuselage and the modified (fig. 10(b)) showed that subsonically the aft model

aft fuselage for the forward model configuration configuration had a larger nosedown pitching moment

(Mach 1.1, 7-ft wind tunnel), increment than the forward model configuration. To

identify the reason for this, A502 code analyses

f,

were completed at Mach 0.6 for the baseline SR-71

canoe-only configuration with the SR-71 aft fuselage airplane, canoe-only, forward model, and aft model

modified for the sting mount, configurations. For the forward configuration, the

A502 code predicted a pitching moment increment of

Figure 15 shows pitchingmomentincrements caused --0.0018; whereas, the wind-tunnel increment was

by the canoe from TranAir and the wind tunnel. The -0.0034. For the aft model configuration, A502 code

TranAir increments agree well with the wind-tunnel predicted an increment of -0.0076; whereas, the winddata. In addition, TranAir validated that the SR-71 aft tunnel increment was -0.0084. Hence, A502 code

fuselage configuration had no major affect on the concurred that the pitching moment became

pitching moment increment, increasingly negative as the model was moved aft.

13

AmericanInstituteof AeronauticsandAstronautics
Cp= 0.5

Cp= -1.0 960246

(a) Baseline SR-7 1 airplane.

Higher pressure

B.L.

Lower pressure

(b) Canoe-only configuration.

Figure 16. TranAir surface pressure maps at Math 1.1 and 4° angle of attack.

14

American Institute of Aeronautics and Astronautics
region forward of the canoe. More significantly, a

.3 -- -O-Baseline SR-71 decreased pressure region extends from the canoe to the

_ga_c_ a_c_y_on. wing at the location where the canoe width starts to

.1 _ decrease. The combination of these effects would cause

0 a nosedown pitching moment for the canoe-only

-.1 configuration. The A502 code predicted a pitching

Op-_ moment increment of-0.0032 for the canoe-only

configuration.

-.3

Momentreference "I:]C] q With the model installed, low-pressure regions are

-.4 -- (F.S.900) h

-.5 -- !_._ _ Canoe _! thseene SoR-7 n th1e leading fuselage edanged ofwitnheg aeros adjacpeike nt mtoodthele anmdodonel

-.6- i I [ [ [ l I'-[ [ (figs. 18(c)and 18(d)). The extremely low-pressnre

-.7 region on top of the model is unrealistic and related to 600 700 800 900 1000 1100 1200 1300

the fact that A502 is a potential flow code. Potential flow Fuselage station, in. s_o2_

codes do not accurately handle separated flow. Instead,

Figure 17. TranAir pressure distributions at B.L. 19.4 such codes accelerate flows around sharp comers;

for the canoe-only configuration at Mach 1.1 and 4° however, this influence on the pitching moment

angle of attack, increment is small. The low-pressure region on the

fuselage and wing adjacent to the model has the most

significant effect on pitching moment. In the forward

The baseline SR-71 airplane, canoe-only, forward model position, this low-pressure region is forward of

model, and aft model configuration pressure maps are the moment reference location (F.S. 900) and, hence,

shown in figures 18(a) through 18(d). Figure 19 shows reduces the nosedown pitching moment. For the aft

the chordwise pressure distributions for the 22-percent model configuration, this low-pressure region combines

semispan wing station. Note that for the canoe-only with the low-pressure flow around the aft end of the

configuration, comparing figures 18(a) with figure 18(b) canoe to increase the nosedown pitching moment.

and from figure 19 reveals a slightly increased pressure

Cp=0.5

Cp= -1.0 t_oa4o

(a) Baseline SR-71 airplane.

Figure 18. A502 surface pressure maps at Mach 0.6 and 4° angle of attack.

15

Aancrican Institute of Aeronautics and Astronautics
Higher

Lower

(b) Canoe-only configuration.

F.S. 900

Low pressure region

adjacent to the model

96O251

(c) Forward model configuration.

Figure 18. Continued.

16

Amcrican Institute of Acronautics and Astronautics
F.S. 900

Low pressure region

adjacent to the

960252

(d) Aft model configuration.

Figure 18. Concluded.

.5 -- -O- Baseline SR-71 . 100 x 103

-O- Canoe-only configuration ._ -- O LASREfligM test points

0 "O"Forward model configuration _" 90 --

-.10 6o

Altitude, 310 K

Cp

ft .50 D# 0 _ _--NormalSR-71

-.20 Moment reference ""_ 403 F _tfoJs00 KEAS f,lghtenv.lope

(F.S. 900) Additionalrestrictions

c.no. ,o

-.s, I I I I I I

700 800 900 1000 1100 1200 1300 0 .S 1.0 1.5 2.0 2.5 3.0 3.S 4.0

Fuselage station, in. _o2s3 Mach number _o_

Figure 19. A502 wing pressure distributions at the Figure 20. The LASRE flight envelope with restricted

22-percent semispan station, region caused by predicted hinge moment limits.

Flight SimulationControlandPerformanceAnalysis SR-71 surfaces are rigged such that the outboard

elevons become hinge moment limited before the

Flight simulations demonstrated that the additional inboards. The LASRE vehicle was restricted to

pitching moment increment for the LASRE airspeeds below that which the outboard elevons would

configuration would reduce airspeed and e.g. operating become hinge moment limited. On the one hand, the

limits as a result of devon hinge moment limits, required LASRE flight test points do not occur in this

Figure 20 shows the normal and modified SR-71 restricted region. On the other hand, the restricted part

airspeed envelopes. The SR-71 control surfaces are of the envelope was no longer available for the transonic

actuated by two hydraulic systems. Safety of flight acceleration which results in a less fuel efficient

constraints require that the aircraft be controllable in l-g acceleration.

flight with only one hydraulic system operational. The

17

American Institute of Aeronautics and Astronautics
Figure 21 shows how the airspeed limits were derived. 70 x 10a

These limits were based on the outboard hinge moment I-- -- Levelacceleration,

being limited in a dive maneuver. In this scenario, the 1 no enhancement

vehicle climbs to an altitude of 33,000 ft at a constant 65 - - - Level acceleration,

with enhancement

Mach 0.9 followed by a constant rate dive. Constant 60

descent rates from 1000 to 6000 ft/min are plotted in --- Climb-diveacceleration, with enhancement figure 21 for a 120,000 lbf aircraft with the e.g. at 55

24-percent m.a.c. The outboard actuators are hinge

moment limited at equivalent airspeeds of 50

approximately 465 to 475 kn. The inboard elevens do _

not become limited for at least another 25 kn. Fuel 45

quantity,

The piloted, real-time simulation was used with the Ib 40

LASRE mass and drag properties to estimate the

maximum Mach number attainable with the aft model 35 __"x

configuration. The simulation was run with standard 30 "_

J-58 engines (Pratt & Whitney, Palm Beach, Florida) _

and with engines upgraded for enhanced thrust 25 _

performance. The majority of the thrust enhancement Return to base---_ x_

was derived from manually uptrimming the exhaust gas 20

temperatures from Mach 1.0 to Mach 2.5. The entire I I I I

pod weight was assumed to be 13,838 lb. 15 1.0 1.5 2.0 2.5 3.0

Figure 22 shows a level acceleration at an altitude of Machnumber 96o256

25,000 ft from Mach 0.75 to 450 KEAS followed by a

450 KEAS climb at standard day atmospheric Figure 22. Performance simulation results for the

conditions. For operational reasons, the aircraft LASRE configuration with and without thrust

was required to cease testing and return to base when enhancement.

20,000 lb of fuel remained. This requirement includes

5000 lb of unusable fuel which was needed as forward

ballast to counter the additional aft-mounted weight of the aerospike pod. With thrust enhancement, the vehicle

reached the desired Mach 3.2 with enough fuel

remaining for approximately 2.5 min of cruise flight

before having to return to base. Without enhancement

Outboard hinge moment limits

withone hydraulic system the vehicle only reached Mach 3.09 before having to

BZ_ Inboardhingemomentlimits return to base.

60x1__3_0 withonehydraulicsystem

Figure 22 also shows results using a transonic climb

55 _-- and dive maneuver. Starting at Mach 0.75 and an

5o _s altitude of 25,000 ft, the vehicle accelerated at level

as altitude to Mach 0.9 for a constant Mach number climb

40 to an altitude of 33,000 ft. At this point, a 3000 ft/min

Altitude, descent was begun. At an altitude of 26,000 ft, a 450 

It 35

oo KEAS climb was initiated. Using thrust enhancement,

30 . KEAS= 4_ 1000fpm the vehicle accelerated to Mach 3.2 with enough fuel

25 ---- KEAS= 47_^" 200Cfpm2000 remaining for 3.6 min of cruising flight before return to

..... _]_I,_000 3000fpm base. Hence, the climb and dive maneuver provided a

2015 -- _fpm small performance benefit in terms of fuel saved during

10 ] ] ] I I ] ] ] the transonic portion of the acceleration.

.8 .9 1.0 1.1 1.2 1.3 1.4 1.5 1.6

Mach number Concluding Remarks

960255

Figure 21. Airspeed limits in a dive maneuver caused A series of wind-tunnel tests was completed in

by predicted outboard and inboard eleven hinge support of the Linear Aerospike SR-71 Experiment

moment limits.

18

American Institute of Aeronautics and Astronautics
flight test program. These tests identified problems with These simulations focused on control authority and

the initial forward model configuration caused by an performance. Envelope restrictions were applied to the

unexpected and excessive nosedown pitching moment. Linear Aerospike SR-71 Experiment configuration

In addition, an aft model configuration that is expected based on hinge moment limits. The real-time simulation

to meet performance, stability, and control requirements was used to verify that the aircraft could be accelerated

during the flight test program was identified. The cause through the transonic Mach numbers. In addition, the

of the nosedown pitching moment was investigated by maximum desired Mach number of Mach 3.2 was

component buildup in the wind tunnel and with obtainable although only a few minutes of test time

potential flow computational analyses. For supersonic were available before having to cease testing and return

Mach numbers, the nosedown pitching moment was to base.

caused by the combination of a high-pressure region

forward of the canoe because of the bow shock and a APPENDIX A

,, low-pressure region aft of the canoe because of the

expansion fans. FORWARD MODEL

AERODYNAMIC INCREMENTS

The canoe alone caused a significant nosedown

pitching moment. Adding the model in the forward

position increased the nosedown pitching moment; This appendix describes the aerodynamic model

whereas, adding the model in the aft position negated increments for the forward LASRE configuration

the majority of the canoe-induced nosedown pitching (fig. 4). In creating these tables, conservative (read

moment. The additional trim drag and elevon actuator pessimistic) engineering judgements were always made

limits associated with the forward model configuration in regard to LASRE performance, stability, and control.

prevented the vehicle from being able to accelerate Data for this configuration are from the 4-ft wind tunnel

through the transonic speeds and, hence, made this for Mach 0.6 through Mach 3.2. A complete

configuration unacceptable. At subsonic Mach numbers, aerodynamic model was not obtained in the wind tunnel

the nosedown pitching moment was worse for the aft because this configuration was not acceptable for the

model configuration. Potential flow analysis showed LASRE experiment. The data that were obtained are

that this effect was caused by a low-pressure region presented here to document the aerodynamic

created on the fuselage and wing adjacent to the model, characteristics of a payload mounted mostly forward of

The additional subsonic trim drag for the aft model the pitching moment reference on top of the SR-71

configuration did not significantly affect performance; airplane.

hence, the aft model configuration was shown to have

acceptable drag and stability increments. Tables A-1 through A-3 list the drag, lift, and pitching

moment increments as a function of Mach number and

Ground-based flight simulations were critical in wing reference plane angle of attack.

assessing the acceptability of these configurations.

Table A- 1. The CD increment caused by the LASRE pod with the model in the forward position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0 ° 1° 2° 3° 4° 5° 6° 7 ° 8°

0.90 0.0039 0.0046 0.0050 0.0053 0.0054 0.0058 0.0059 0.0065 0.0066 0.0062 0.0082

0.95 0.0062 0.0059 0.0054 0.0086 0.0090 0.0080 0.0100 0.0095 0.0104 0.0093 0.0104

1.05 0.0082 0.0085 0.0089 0.0101 0.0109 0.0113 0.0121 0.0128 0.0128 0.0134 0.0126

1.10 0.0088 0.0090 0.0103 0.0109 0.0115 0.0124 0.0128 0.0133 0.0133 0.0133 0.0123

1.20 0.0082 0.0086 0.0094 0.0100 0.0105 0.0111 0.0113 0.0113 0.0114 0.0106 0.0095

1.40 0.0066 0.0070 0.0074 0.0079 0.0077 0.0070 0.0076 0.0075 0.0068 0.0071 0.0066

1.60 0.0058 0.0060 0.0065 0.0070 0.0068 0.0066 0.0068 0.0064 0.0060 0.0055 0.0055

2.00 0.0055 0.0060 0.0058 0.0058 0.0055 0.0051 0.0051 0.0052 0.0048 0.0045 0.0042

19

AmericanIzlstituteof AeronauticsandAstronautics
Table A-2. The CL increment caused by the LASRE pod with the model in the forward position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0° 1° 2° 3° 4 ° 5° 6° 7 ° 8 °

0.90 0.0090 0.0080 0.0090 0.0090 0.0110 0.0090 0.0060 0.0080 0.0060 0.0060 0.0110

0.95 0.0020 0.0130 0.0180 0.0070 0.0070 0.0130 0.0110 0.0070 0.0060 0.0080 0.0070

1.05 0.0150 0.0170 0.0160 0.0200 0.0210 0.0200 0.0220 0.0210 0.0200 0.0150 0.0090

1.10 0.0200 0.0210 0.0210 0.0250 0.0240 0.0260 0.0260 0.0230 0.0230 0.0220 0.0180

1.20 0.0170 0.0150 0.0150 0.0180 0.0170 0.0180 0.0170 0.0140 0.0110 0.0040 0.0000

1.40 0.0070 0.0060 0.0030 0.0040 -0.0020 -0.0040 -0.0020 -0.0010 -0.0060 -0.0020 -0.0060

1.60 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000 0.0000

2.00 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100 -0.0100

Table A-3. The Cm increment caused by the LASRE pod with the model in the forward position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0 ° 1° 2° 3° 4 ° 5° 6° 7 ° 8°

0.90 -0.0039 -0.0034 -0.0039 -0.0044 -0.0041 -0.0042 -0.0042 -0.0059 -0.0068 -0.0069 -0.0111

0.95 -0.0027 -0.0063 -0.0075 -0.0037 -0.0038 -0.0057 -0.0050 -0.0061 -0.0064 -0.0073 -0.0092

1.05 -0.0072 -0.0084 -0.0101 -0.0116 -0.0130 -0.0146 -0.0160 -0.0169 -0.0162 -0.0159 -0.0147

1.10 -0.0077 -0.0115 -0.0129 -0.0143 -0.0160 -0.0183 -0.0208 -0.0219 -0.0222 -0.0227 -0.0225

1.20 -0.0129 -0.0142 -0.0160 -0.0175 -0.0190 -0.0199 -0.0199 -0.0191 -0.0197 -0.0156 -0.0137

1.40 -0.0163 -0.0144 -0.0147 -0.0146 -0.0132 -0.0129 -0.0128 -0.0124 -0.0115 -0.0119 -0.0119

1.60 -0.0160 -0.0140 -0.0130 -0.0135 -0.0130 -0.0120 -0.0105 -0.0095 -0.0090 -0.0095 -0.0095

2.00 -0.0055 -0.0055 -0.0050 -0.0050 -0.0045 -0.0045 -0.0045 -0.0043 -0.0040 -0.0045 -0.0050

Table A-4 lists the side force, rolling moment, and Table A-4. Sideslip derivative increments caused by the

yawing moment sideslip derivative increments. These LASRE pod with the model in the forward position.

data were obtained using sideslip sweeps at a set angle Mach Angle of ACyf_, AClf, AC

of Machattack. The approximate angle of attack for each number attack, deg per deg per deg per de_'g number is listed in the table.

0.60 4.1 ° -0.00138 0.00031 -0.00078

0.95 5.2 ° -0.00055 0.00061 -0.00101

1.20 4.8 ° -0.00051 0.00089 -0.00104

2.00 4.6 ° -0.00008 0.00053 -0.00076

2O

American Institute of Aeronautics and Astronautics
APPENDIX B Because of the relatively small size of the aerospike

pod, however, the viscous drag correction, at most, gives

AFT MODEL AERODYNAMIC a reduction of six drag counts. In addition, these drag

INCREMENTS data do not include trim drag increments associated with the aft model configuration. Wind-tunnel data show that

drag resulting from elevon deflection was not affected

This appendix describes the aerodynamic model by the model in the aft configuration.

increments for the LASRE configuration with the model

mounted in the aft position (fig. 5). In creating these Table B-2 lists the lift increment as a function of

tables, conservative engineering judgements were Mach number and wing reference plane angle of attack.

always made in regard to LASRE performance, stability, These data were obtained from the 7-ft tunnel and from

and control. Data for this configuration are from the 7-ft Mach 0.6 to Mach 1.6. Hold last values were used at

_, wind tunnel for Mach 0.6 through Mach 1.6 and from instances above Mach 1.6 and below Mach 0.6.

the 4-ft wind tunnel for the higher Mach numbers.

Table B-3 lists the pitching moment increment as a

Table B-1 lists the drag increment as a function of function of Mach number and wing reference plane

Mach number and wing reference plane angle of attack, angle of attack.

Drag data at Mach 1.6 were not used because of an

anomaly in these data. Mach 2 data were obtained in the Table B-4 lists the pitching moment elevon

first 4-ft wind tunnel test using the initial aft-mounted effectiveness increment as a function of Mach number.

model configuration. This initial configuration has Only the inboard elevons were adjusted in these tests

similar supersonic drag increments as the final aft- because the outboard elevons were not expected to be

mounted model configuration (fig. 6(a)). The Mach 2.5 affected by the LASRE pod. The elevons used in the 4-ft

and Mach 3.2 data were extrapolated from the Mach 2.0 wind tunnel at Mach 3.2 had a 33-percent span

data using engineering analysis to adjust for the reduction because of the aft end modification

to accommodate the sting mount. The increment changing wave drag. No viscous drag corrections were

made to account for the fact that wind-tunnel Reynolds measured at Mach 3.2 was adjusted for a full-span

numbers were less than flight Reynolds numbers. The elevon; however, no significant increment in elevon

flight viscous drag component would be lower than the effectiveness was measured at Mach 3.2.

wind tunnel because of higher flight Reynolds numbers.

Table B-1. The Co increment caused by the LASRE pod with the model in the aft position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0° 1° 2° 3° 4° 5° 6° 7° 8°

0.00 0.0038 0.0038 0.0040 0.0042 0.0041 0.0041 0.0036 0.0034 0.0032 0.0028 0.0033

0.60 0.0038 0.0038 0.0040 0.0042 0.0041 0.0041 0.0036 0.0034 0.0032 0.0028 0.0033

0.90 0.0102 0.0103 0.0105 0.0105 0.0103 0.0100 0.0095 0.0091 0.0087 0.0082 0.0087

0.95 0.0143 0.0143 0.0143 0.0139 0.0132 0.0133 0.0129 0.0125 0.0117 0.0107 0.0109

1.05 0.0128 0.0125 0.0120 0.0116 0.0111 0.0106 O.OlO1 0.0095 0.0091 0.0088 0.0085

" 1.10 0.0122 0.0116 0.0112 0.0107 0.0101 0.0098 0.0094 0.0089 0.0084 0.0082 0.0081

1.20 0.0097 0.0091 0.0089 0.0085 0.0083 0.0081 0.0078 0.0078 0.0075 0.0075 0.0073

1.40 0.0066 0.0067 0.0066 0.0066 0.0067 0.0066 0.0065 0.0062 0.0058 0.0056 0.0054

1.60 0.0060 0.0061 0.0060 0.0060 0.0059 0.0058 0.0056 0.0055 0.0051 0.0047 0.0045

2.00 0.0049 0.0050 0.0047 0.0048 0.0044 0.0041 0.0040 0.0040 0.0037 0.0029 0.0025

2.50 0.0046 0.0047 0.0044 0.0045 0.0041 0.0038 0.0037 0.0037 0.0034 0.0026 0.0022

3.20 0.0043 0.0044 0.0041 0.0042 0.0038 0.0035 0.0034 0.0034 0.0031 0.0023 0.0019

21

American Institute of Aeronautics and Astronautics
Table B-2. The C L increment caused by the LASRE pod with the model in the a_ POsition.

Angle of attack, deg

Mach

number -2 ° -1 ° 0 ° 1° 2° 3° 4° 5° 6 ° 7 ° 8 °

0.00 0.0046 0.0033 0.0052 0.0073 0.0017 0.0033 -0.0035 -0.0048 -0.0057 -0.0084 -0.0068

0.60 0.0046 0.0033 0.0052 0.0073 0.0017 0.0033 -0.0035 -0.0048 -0.0057 -0.0084 -0.0068

0.90 -0.0010 -0.0052 -0.0063 -0.0094 -0.0100 -0.0112 -0.0144 -0.0149 -0.0165 -0.0172 -0.0133

0.95 -0.0224 -0.0211 -0.0220 -0.0220 -0.0192 -0.0181 -0.0190 -0.0200 -0.0213 -0.0249 -0.0287

1.05 -0.0180 --0.0183 -0.0183 -0.0187 -0.0170 -0.0188 -0.0200 -0.0211 -0.0235 -0.0224 -0.0184

1.10 -0.0190 -0.0193 -0.0164 -0.0168 -0.0175 -0.0160 -0.0168 -0.0176 -0.0188 -0.0185 -0.0216

1.20 -0.0120 -0.0115 -0.0110 -0.0109 -0.0106 -0.0100 -0.0097 -0.0093 -0.0103 -0.0118 -0.0144

1.40 -0.0099 -0.0095 -0.0094 -0.0099 -0.0098 -0.0103 -0.0120 -0.0132 -0.0133 -0.0136 -0.0084

1.60 -0.0069 -0.0072 -0.0072 -0.0074 -0.0074 -0.0061 -0.0068 -0.0071 -0.0081 -0.0077 -0.0081

2.00 -0.0069 -0.0072 -0.0072 -0.0074 -0.0074 -0.0061 -0.0068 -0.0071 -0.0081 -0.0077 -0.0081

2.50 -0.0069 -0.0072 -0.0072 -0.0074 -0.0074 -0.0061 -0.0068 -0.0071 -0.0081 -0.0077 -0.0081

3.20 -0.0069 -0.0072 -0.0072 -0.0074 -0.0074 -0.0061 -0.0068 -0.0071 -0.0081 -0.0077 -0.0081

Table B-3. The Cm increment caused by the LASRE pod with the model in the a_ position.

Angle of _tack, deg

Mach

number -2 ° -1 ° 0° 1° 2 ° 3° 4 ° 5° 6° 7° 8°

0.00 -0.0086 -0.0083 -0.0085 -0.0090 -0.0087 -0.0091 -0.0084 -0.0083 -0.0079 -0.0078 -0.0074

0.60 -0.0086 -0.0083 -0.0085 -0.0090 -0.0087 -0.0091 -0.0084 -0.0083 -0.0079 -0.0078 -0.0074

0.90 -0.0105 -0.0090 -0.0090 -0.0087 -0.0093 -0.0098 -0.0092 -0.0094 -0.0091 -0.0087 -0.0071

0.95 0.0038 0.0029 0.0027 0.0026 0.0010 -0.0002 -0.0006 -0.0008 -0.0008 0.0000 0.0010

1.05 0.0032 0.0036 0.0031 0.0040 0.0039 0.0039 0.0040 0.0040 0.0041 0.0032 0.0032

1.I0 0.0030 0.0034 0.0027 0.0026 0.0027 0.0027 0.0020 0.0015 0.0013 0.0006 0.0015

1.20 -0.0013 -0.0015 -0.0022 -0.0025 -0.0027 -0.0030 -0.0033 -0.0035 -0.0036 -0.0032 -0.0026

1.40 -0.0017 -0.0015 -0.0016 -0.0014 -0.0012 -0.0011 -0.0008 -0.0006 -0.0011 -0.0013 -0.0019

1.60 -0.0011 -0.0011 -0.0011 -0.0010 -0.0008 -0.0010 -0.0010 -0.0010 -0.0007 -0.0005 -0.0003

2.00 0.0000 0.0002 0.0001 0.0000 0.0000 0.0000 0.0000 0.0002 0.0005 0.0011 0.0014

2.50 0.0016 0.0014 0.0008 0.0009 0.0011 0.0011 0.0014 0.0010 0.0008 0.0007 0.0011

3.20 0.0029 0.0026 0.0024 0.0027 0.0024 0.0023 0.0021 0.0022 0.0023 0.0022 0.0024

Table B-4. Elevon Cm effectiveness Tables B-5 through B-7 list the side force, rolling

increment caused by the LASRE pod moment, and yawing moment sideslip derivative

with the model in the aft position, increments.

Mach ACm_e,

number per deg Table B-8 lists the side force, rolling moment, and

0.0 0.00017 yawing moment rudder control derivative increments as

a function of Mach number. 0.6 0.00017

1.1 0.00028

1.4 0.00008

3.2 0.00000

22

AmericanInstituteof AeronauticsandAstronautics
Table B-5. The Cy[_ increment (per deg) caused by the LASRE pod with the model in the aft position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0° 1° 2° 3° 4 ° 5° 6° 7° 8°

0.00 -0.00092 --0.00070 -0.00080 -0.00127 -0.00064 -0.00112 -0.00106 -0.00114 -0.00120 -0.00110 -0.00114

0.60 -0.00092 -0.00070 -0.00080 -0.00127 -0.00064 -0.00112 -0.00106 -0.00114 -0.00120 -0.00110 -0.00114

0.95 -0.00144 --0.00160 -0.00133 -0.00145 -0.00142 -0.00154 -0.00127 -0.00082 -0.00073 -0.00070 -0.00040

1.10 -0.00027-0.00007 0.00012 0.0(X)26 0.00045 0.00054 0.00067 0.00022 0.00083 0.00032 0.00105

1.60 -0.00215 -0.00236-0.00233-0.00225-0.00219-0.00212-0.00241 -0.00208 -0.00216 -0.00227 -0.00199

i, 2.00 -0.00095 -0.00116 -0.00113 -0.00105 -0.00099 -0.00092 -0.00121 -0.00088 -0.00096 -0.00107 -0.00079

3.20 -0.0(X)80 -0.00077 -0.00052 -0.00037 -0.00044 -0.00036 -0.0(X)34 -0.00036 -0.00030 -0.00027 -0.00012

Table B-6. The C1_ increment (per deg) caused by the LASRE pod with the model in the aft position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0 ° 1o 2° 3 ° 4_ 5 ° 6° 7 ° 8°

0.00 0.00023 0.00023 0.00023 0.00022 0.00023 0.00027 0.00025 0.00031 0.00023 0.00037 0.00034

0.60 0.00023 0.00023 0.00023 0.00022 0.00023 0.00027 0.00025 0.00031 0.00023 0.00037 0.00034

0.95 0.00028 0.00024 0.00026 0.00028 0.00032 0.00035 0.00031 0.00035 0.00031 0.00040 0.00048

1.10 0.00051 0.00048 0.00043 0.00044 0.00038 0.00034 0.00025 0.00019 0.00020 0.00019 0.00038

1.60 0.00026 0.00030 0.00030 0.00030 0.00032 0.00031 0.00029 0.00032 0.00031 0.00030 0.00032

2.00 0.00021 0.00025 0.00025 0.00024 0.00026 0.00026 0.00023 0.00027 0.00025 0.00024 0.00026

3.20 0.00010 0.00010 0.00010 0.00009 0.00009 0.00009 0.00007 0.00007 0.00008 0.00008 0.00009

Table B-7. The Cn[_ increment (per deg) caused by the LASRE pod with the model in the aft position.

Angle of attack, deg

Mach

number -2 ° -1 ° 0° 1° 2° 3° 4 ° 5° 6° 7° 8°

0.00 -0.00015 -0.00016 -0.00014 -0.00020 -0.00018 -0.00019 -0.00018 -0.00016 -0.00014 -0.00015 -0.00020

0.60 -0.00015 -0.00016 -0.00014 -0.00020 -0.00018 -0.00019 -0.00018 -0.0(X)16 -0.0(X)14 -0.00015 -0.0(X)20

0.95 0.00018 0.00016 0.00012 0.0(X)13 0.00012 0.00014 0.0(X)08 -0.00(X)5 -0.0(X) 11 -0.0(X)11 -0.0(X) 14

1.10 -0.00038 -0.00052 -0.00064 -0.00073 -0.00076 -0.00078 -0.00078 -0.00078 -0.00075 -0.00071 -0.00079

1.60 0.00029 0.00035 0.00038 0.00041 0.00045 0.00049 0.00051 0.00052 0.00055 0.00059 0.00058

2.00 -0.00003 0.00(X)3 0.O(XX)6 0.00009 0.00013 O.O(X)17 0.00019 0.0(X)20 0.00023 0.00027 0.00026

3.20 -0.00015 -0.00017 -0.00021 -0.00023 -0.00019 -0.00020 -0.00016 -0.00011 -0.00012 -0.00010 -0.00011

23

American Institute of Aeronautics and Astronautics

......... -- -- --r--r •......
Table B-8. Rudder control derivative increments caused Corporation, Report no. SP-508, Burbank, California,

by the LASRE pod with the model in the aft position. Oct. 29, 1964.

4Saaris, Gary R., A5021 User's Manual - PAN AIR

Mach ACy8 r, meier' ACn_r' Technology Program for Solving Problems of Potential

number per deg per deg per deg Flow about Arbitrary. Configurations, Document no.

0.0 -0.0001148 0.0000552 0.0000597 D6-54703, The Boeing Company, Seattle, Washington,

Feb. 1992.

0.6 -0.0001148 0.0000552 0.0000597

1.1 -0.0006052 0.0000409 0.0003059 5Johnson, ET., Samant, S.S., Bieterman, M.B.,

1.6 0.0000000 0.0000238 0.0000000 Melvin, R.G., Young, D.P., Bussoletti, J.E., and Hilmes,

C.L., TranAir: A Full-Potential, Solution-Adaptive,

3.2 0.0000000 0.0000000 0.0000000

Rectangular Grid Code for Predicting Subsonic,

Transonic, and Supersonic Flows About Arbitrao,

Configurations--Theo_ Document, NASA CR 4348,

1992.

References

6johnson, ET., Samant, S.S., Bieterman, M.B.,

IMartinez, A., Aerodynamic Nozzle Study, vol. 1, Melvin, R.G., Young, D.P., Bussoletti, J.E., and Hilmes,

Rocketdyne Final Report R-6582, North American C.L., TranAir: A Full-Potential Solution-Adaptive,

Aviation, Inc., Los Angeles, California, July 15, 1966. Rectangular Grid Code for Predicting Subsonic,

Transonic, and Supersonic Flows About Arbitrary

2Martinez, A., Aerodynamic Nozzle Study, Slipstream Configurations--User's Manual, NASA CR 4349, 1992.

Studies, vol. III, Rocketdyne Interim Report R-6273,

North American Aviation, Inc., Los Angeles, California, 7Meyer, J.E. and Campbell, D.H., The YF-12A as a

July31, 1965. Proto_.pe Inlet Flight Test Bed, Lockheed Aircraft

Corporation - Advanced Development Projects, Report

3Meyer, J.E., McMaster, J.R., and Moody, R.L., no. SP-4137A, Burbank, California, Mar. 15, 1974.

Handling Qualities of the SR-71, Lockheed Aircraft

24

AmericanInstituteof AeronauticsandAstronautics
l--
REPORT DOCUMENTATIONPAGE FormApproved OMB No. 0704-0188

Publicreportingburdenforthiscollectionofinformationisestimatedtoaverage1 hourperresponse,includingthetimeforreviewinginstructions, searchingexistingdatasources,

gatheringandmaintainingthedataneeded,endcompletingandreviewingthecollectionofinformation. Sendcommentsregardingthisburdenestimateoranyotheraspectofthiscollectionofinformation, includingsuggestionsforreducingthisburden,toWashingtonHeadquartersServices,Directorafor te InformationOperationsandReports1215 , JeffersonDavis

Highway,Suite1204,Arlington, VA22202-4302,endtotheOfficeofManagement andBudget,Paper'_vork ReductionProject(0704-0188), WashingtonDC, 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORTTYPE AND DATES COVERED

June 1996 TechnicalMemorandum

4.TITLE AND SUBTITLE 5. FUNDING NUMBERS

Wind-Tunnel Development of an SR-71 Aerospike Rocket Flight Test

Configuration

6. AUTHOR(S) WU 505-68-52

TimothyR. Moes,BrentR.Cobleigh,TimothyR.Connors,TimothyH.

Cox, Stephen C. Smith, and Norm Shirakata

7.PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8.PERFORMING ORGANIZATION

REPORT NUMBER

NASA Dryden Flight Research Center

P.O. Box 273 H-2108

Edwards, California 93523-0273

9. SPONSORING/MONOTORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

National Aeronautics and Space Administration

Washington, NASA TM-4749 DC20546-0001

11.SUPPLEMENTARYNOTES

Presented as AIAA 96-2409 at the 14th Applied Aerodynamics Conference, New Orleans, Louisiana, June 17-20, 1996. Timothy R.

Moes, Brent R. Cobleigh, Timothy R. Conners, and Timothy H. Cox, NASA Dryden Flight Research Center, Edwards, CA; Stephen C.

Smith, NASA Ames Research Center, Moffett Field, CA; Norm Shirakata, Lockheed-Martin Skunk Works, Palmdale, CA.

12a.DISTRIBUTION/AVAILABILITY STATEMENT 12b.DISTRIBUTIONCODE

Unclassified--Unlimited

Subject Category 02

13. ABSTRACT (Maximum 200 words)

A flight experiment has been proposed to investigate the performance of an aerospike rocket motor installed in a lifting

body configuration. An SR-71 airplane would be used to carry the aerospike configuration to the desired flight test conditions. Wind-tunnel tests were completed on a 4-percent scale SR-71 airplane with the aerospike pod mounted in various

locations on the upper fuselage. Testing was accomplished using sting and blade mounts from Mach 0.6 to Mach 3.2. Initial test objectives included assessing transonic drag and supersonic lateral--directionalstability and control. During these

tests, flight simulations were run with wind-tunnel data to assessthe acceptability of the configurations. Early testing demonstrated that the initial configuration with the aerospike pod near the SR-71 center of gravity was unsuitable because of

large nosedown pitching moments at transonic speeds. The excessive trim drag resulting from accommodating this pitching moment far exceeded the excessthrust capability of the airplane. Wind-tunnel testing continued in an attempt to find a

configuration suitable for flight test. Multiple configurations were tested. Results indicate that an aft-mounted model configuration possessed acceptable performance, stability, and control characteristics.

14. SUBJECTTERMS 15. NUMBER OF PAGES

Aerodynamic models; Aerospike; Flight simulations; Rocket; SR-71 airplane; Wind 27

tunnel 18.PRICECODE A03

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATIONOF ABSTRACT

OF REPORT OFTHIS PAGE OF ABSTRACT

Unclassified Unclassified Unclassified Unlimited

NSN 7540-01-280-5500 Available from the NASA Center forAeroSpace Information, 800 Elkridge Landing Road, Standard Form298 (Rev.2-89)

Linthicum Heights, MD 21090; (J01)621-0390 Presc 298-102 ribbyedANSStdI Z3. 9-18

3 1176 01426 1979

National Aeronautics and SPECIALFOURTH-CLASS RATE

Space Administration POSTAGEANDFEESPAID NASA

Code JTT PERMITNoG27

Washington,D.C.20546-0001

USA

OfficialBusiness

Penaltyfor PrivateUse,$300

POSTMASTER:IfUndeliverabl(Se ection158

Postamanua l l)Do NotReturn