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AlAA 93-0036 Wind Tunnel Investigation of a Missile with an Operational Midbody Turbojet Engine

Lamar M. Auman U.S. Army Missile Command Redstone Arsenal, Alabama

31st Aerospacb Sciences Meeting & Exhibit

January 11 -1 4, 1993 / Reno, NV For permission to copy or republish, contact the American institute of Aeronautics and Astronalrtics 370 L'Enfant Promenade, S.W., Washington, D.C. 20024 :

WIND TUNNEL INVESTIGATION WITH AN OPERATIONAL TURBOJET ENGINE

Lamar M. Auman* US h y Missile Command Redstone Arsenal, Alabama

ABSTRACT A wind tunnel test of a turbojet powered

missile with the engine installed and operating was conducted in Calspan's Transonic Wind Tunnel in Buffalo, New York from 20 May to 3 June 1991. The primary model corrfguration consisted of a bifurcated inlet and exhaust which was sting mounted. Insmmentation consisted of a six-component main balance, four fin balances, 39 thermocouples, 10 heat flow gages and several base pressure ports. The purpose of this test was to investigate (1) stability and control, (2) installed engine performance at flight conditions, and (3) plume interactions or aft heating problems. The investigation obtained data at Mach numbers from 0.05 to 0.80, with the majority of the data collected at Mach 0.33, 0.50 and 0.60. Angle of attack was swept from - 1 6 O to +16', with engine throttle setting, fin deflection angle and Mach number held constant. Missile stability data indicates that the operational mid-body turbojet engine has no adverse effect on missile stability or control authority. Aftbody thermal data indicates that the worst case occurs during a low Mach number (static) high kngine throttle setting condition.

NOMENCLATURE

L,

CAoff Equivalent engine-off m i d force

C A ~ , Engine-on axial force coefficient. CD Wind axis drag coefficient. CM Pitching moment coefficient. C, Normal force coefficient. Cy Side force coefficient.

coefficient.

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* Aerospace Engineer llk pper is a declared work of the U.S. Government and is not subject to copyright protection in the United States.

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CF,, Net system thrust coefficient. d Reference Length (7.00-in). DA

FEX

Airframe drag: positive in the direction of forward flight. Total axial force seen by the system; positive in the direction of forward flight. Installed engine thrust, including all throttle dependent drag forces such as inlet spillage drag, exhaust jet drag, and auxiliary aimow drag; positive in the direction of forward flight.

Fpp

HF Heat Flux gage. kRPM Engine RPM (in thousands). WMcCorrected Engine FWM (in thousands). MRP Moment Reference Point, MS 40.39. MS Model Station (inches). S Reference area, (38.49-in2).

CL Angle of Attack. p Side Slip Angle.

@ Roll angle. Fin deflection angle.

INTRODUCTION Expendable, low-cost turbojet engines

offer attractive alternatives as an efficient sustainer propulsion system for a number of tactical missile systems. Advantages inherent to turbojet engines are (1) variable thrust, (2) high fuel efficiency, (3) minimum visible and infrared signature and (4) compliance with insensitive munitions requirements. Traditional installation has located the turbojet engine in the extreme aft section of the missile with a singular flush mounted inlet and an axial exhaust. However, design requireinents of several future missile systems call for guidance and control equipment such as laser receivers, fiber optic bobbins and/or wire bobbins to be installed in the missile base. This

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requirement forces the propulsion system to accommodate a mid-airframe installation and

I Figure 1. Power-on wind tunnel model installed in I the Calspan test section. ~ ~~~ ~ ~

Proper design and simulation of a turbojet powered bifurcated inlet/exhaust missile must address several design concerns not found with conventional missile systems. Among these concerns are (1) how will the exhaust plume affect missile stability and control authority, ( 2 ) what is the thermal impact of the exhaust plume on the missile afterbody and (3) how will the engine perform in this type of installation.

To address these engineering concerns, a wind tunnel test program was designed to obtain aerodynamic, thermal and engine data on an operational mid-airframe installed turbojet engine. The test was conducted in Calspan’s Transonic Wind Tunnel in Buffalo, New York from 20 May to 3 June 1991. A photograph of the model installed in Calspan’s tunnel is presented in Figure 1. The sting mounted model was pitched in angle of attack of 516’ and a roll pod provided remote roll capability.

EXPERIMENTAL SETUP Wind Tunnel Model The baseline model configuration was a 7

inch diameter 79.58 inch length airframe with four wings, fins and actuator housings in-line, and a bottom mounted wiring race. The missile

was designed to fly in a cruciform configuration with side mounted inlets and a bottom mounted wiring race. To achieve a 5x safety factor and pass the required fuel and

engine shell was designed-and the wiring race was modified. Engine Model Section and Controller

Two separate engine shells (externally identical) were designed and fabricated for the Williams WJ119 engine and the Sundstrand TJ-90 engine. The sections were connected to the model at common forward and aft bulkheads. The primary purpose of the engine shell was to remove the engine from the load path to achieve the 5x safety factor required by Calspan. The diameter of the engine shell was increased to a 7.938 diameter in the region in front of the exhaust. The leading ramp had an angle of 4.00 degrees and began at missile station 28.835. However, the area in front of the inlet was held at a diameter of 7.00 inches and blended into the leading ramp. This was done to insure correct boundary layer height at the inlet face. The trailing ramp had an angle of 7.00 degrees and ended at station 52.946. By thickening the fuselage, the engine was removed from the load path and the wings were attached to the engine shell. The missile wiring race was also modified to allow the engine control Iines and high pressure crank air and start air tubing to be passed around the 7 inch diameter engine.

Both engines were monitored and controlled by MICOM’s Generic Developmental Turbojet Controller(’). This system consists of personal computer equipped with commercially available A D , D/A, frequency converters and several other peripherals.

For the purposes of this investigation, the controller was set up to receive a “taking data” signal from the Calspan data acquisition system. This signal was used start the engine and initialize the two clocks so that the data

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instrumentations lines around the engine, an . .

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could be correlated later. The engine service panel was used to

regulate and control (solenoids controlled by the engine controller) the crank air, torch air, torch hydrogen, JP-10 fuel, and the “Quick- disconnect” mechanism. The panel also served as a “patch panel” for all engine related instrumentation. Williams Engine

The two Williams engines (serial number 301 and 302) built for the wind tunnel test were modifications of the WJ119-2. The WJ119-2, presented in Figure 2, is an axial

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I I diameter turbojet engine.

compressor, axial flow annular combustor, and a single stage cast axial turbine. The inlets and exhaust ducts were integral with the engine and -were a bifurcated design. Several modifications were made to the engines to meet model interfaces and to provide longer engine run capability. The key changes to the engine were the mounting features, and the bearing oil mist system.

The engine was mounted into a clamshell section and the wing mount lugs were removed. At the forward engine bulkhead, 12 threaded holes were used to secure the engine into the model. The WJ119-2 had a self- contained fuel system which was replaced by a new fuel pump, dome regulator, and an external fuel line feed. The pyrotechnic start

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cartridge was replaced by a high pressure air impingement starter and the pyrotechnic ignition torch was replaced by a hydrogenlair ignitor torch.

The baseline engine bearings were modified to incorporate an oil mist generator. The modified lubrication system allowed the engine to operate up to one hour before it became necessary to refill the lubrication system.

The WJ119-2 turbine was designed for a maximum RPM of 58,500 and was spin tested by the manufacturer to 60,000 RPM. After the spin test, the turbine was “Zyglo penetrate” tested for fractures, and the turbine tip diameter was inspected. The maximum operation RPM in the wind tunnel investigation was 56,000, and the controller “auto-kill” RPM was 57,500. Sundstrand Eneine

The TJ90 is a 6.6 inch diameter, 100 lb. thrust class, simple cycle turbojet engine. Air entered the engine axisymmetric inlet housing, after passing through a bifurcated missile intake. The inlet connected to a high specific speed, single stage, backswept radial compressor. The compressor impeller was cast back-to-back with the single stage radial inflow turbine rotor to form the so-called “monorotor”. The support shaft and two angular contact bearings were housed in the inlet structure. Maximum design rotor speed was 104,000 RPM and the maximum tested rotor speed was 102,000 RPM. Air discharging the compressor diffuser passed into a reverse flow annular combustor where the JPlO fuel was burned, then into a radial turbine nozzle. The turbine nozzle and compressor diffuser were formed from a single piece casting or “monostator”. The turbine exhaust gas flowed into a bifurcated exhaust nozzle, a part of the combustor housing assembly. Exhaust gas total temperature was nominally 1650’ F.

Fuel pressure was provided by a centrifugal fuel pump integrally machined into

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the rotor shaft. A carbon seal prevented leakage into the bearings. For long life operation (i.e. wind tunnel testing), the bearings were lubricated by aidoil mist. For short life tactical application, the JPlO fuel would be used as the lubricant. The flow of fuel from the pump was metered by a servo control mounted on the engine air inlet.

Starting was accomplished by high pressure air mounted between the two exhaust duct branches. A single transfer tube conducted the high pressure aidgas to two impingement nozzles directed onto the turbine blades. The JPlO fuel ignition source is a single electrical spark igniter or pyrotechnic flare inserted into the back of the combustor dome.

The Sundstrand engine serial number 23 was the primary Sundstrand engine, and serial number 16 was the backup. Liftin? Surfaces

The wings were NACA 0015 airfoils with a 6-in chord and a 24.5-in. exposed semi-span and mounted directly to the engine shell at missile station 41.83.

The tail fms were NACA 0015 airfoils with a 3.0- in. chord, a 7.5-in. exposed semi- span, and mounted at missile station 73.63. The fins and fin mounts were designed to be compatible with MICOM’s fin balance adapter. The fin deflections were set manually and could be deflected &30° in 2’ increments. Figure 3 shows the location of the fin thermal instrumentation and Figure 4 shows a close-up view of the fin root and thermal instrumentation. Quick Disconnect Device

In order to reduce the number of pressurized lines crossing the balance, a mechanism was designed (using off-the-shelf AN fittings) that could be disconnected remotely after the engine had successfully ignited, This mechanism is presented in Figure

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I Figure 3. Fin 1 thermal instrumentation.

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I Figure 4. Fin 1 root close-up.

I Figure 5. “Quick-disconnect” mechanism.

5 and passed high pressure crank air, tourch air and tourch hydrogen to the model.

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Instrumentation 4 A six-component internal strain gage

balance (Task 2.0 inch MK-33) was used to obtain model force and moment data, and fin balances (one six-component and three three- component) were used to obtain body fixed fin force and moment data. In addition to the force and moment instrumentation, ten type-T, fifteen type-K, fourteen type-E thermocouples and ten heat-flow-gages were installed on and in the airframe to obtain afterbody thermal data. Several static pressure ports were installed in the missile base and balance cavity. The engine was instrumented with various thermocouples and pressure transducers to monitor, control and record engine healthand performance.

The Calspan data acquisition system recorded data from the main balance, four fin balances, the afterbody thermal instrumentation and the base and cavity pressures. All engine data were recorded on MICOM's Generic Developmental Turbojet Controller('). The two data acquisition systems were time synchronized so that the two data files could be merged after testing was completed.

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- Data Reduction and ReDeatability All force and moment coefficient data,

excluding the drag coefficient, are presented in the body-fixed axis system where body normal force'passed between the number 1 and 4 wings. To account for the effect of tunnel static temperature, the engine speed is presented in terms of corrected RPM as defined in equation l...

(1) kRPM ambient Temperature deg. R

518.67

kRPMc =

All drag components were categorized as either airframe drag or throttle dependent drag (or installed net propulsive thrust).

For this test and thrust-drag accounting method, the airframe drag component was assumed to be equivalent to the engine-off drag, and encompassed the external drag associated with (1)- the airframe at the aerodynamic reference condition, and (2) the fin deflections required to trim the airframe. Other drags, such as (1) spillage drag, and (2) jet-effects drag (which are throttle-dependent drag) were assigned to the engine-on drag data. Using these definitions and the discussion found in Reference 2 the thrust/drag accounting system was defined as ...

For the purposes of this test, FEX is assumed to be the engine-on axial force and DA is assumed to be the equivalent engine-off axial force. Nondimensionalizing equation 2 and transforming the terms into the balance axis system yields equation 3.

The equivalent engine-off axial force was calculated for all engine-on runs by applying a third order polynomial curve fit to appropriate engine-off runs. It should be noted that the engine was allowed to windmill while the power-off data was collected. The windmill W M was not recorded but was observed to range around 15 to 20 thousand for several cases, and no provisions were made to calculate air flow through the windmilling engine. Data accuracy and repeatability are summarized in Table 1.

In order to properly assess the engine thrust performance, a detailed thrust-drag accounting method was adopted and rigorously followed.

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Table 1: Main Balance Data Accuracy

~~

Axial Force Ob.)

i l Absolute Accuracy

. .

. .

400.0 5.0% 20.0 0.2% 0.8

RESULTS race. From the data presented in Figure 6 and summarized in Table 2, it is seen that the fin effectiveness is only slightly degraded by the model modifications,

Analysis of the engine-on data, presented in Figures 7 and 8, indicates that engine throttle setting has no effect on missile stability or

Stability and Control Comparisons of engine-off data were made

with data collected during an earlier investigation of the baseline airframe to determine if the airframe modifications affected the basic aerodynamics. Figure 6 presents the pitching moment coefficient versus the normal force coefficient for the

3wer-on entry and the baseline model. The

am .,am .%I) .%W .*.m .2m 0.m %W L* Lrn krn <om

NORMAL FORCE COEFFICIENT. CN

Figure 6. Longitudinal stability comparison for several fin deflections fiom two different wind tunnel enuies.

mtrol authority at Mach 0.33 or 0.60. The dai

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I.@

0.m

.Irn

! 4o.w

1J.W 46.W .?OW 4.w

NORMAL FORCE COEFFICIENT, CN

Figure 7. Longitudinal stability comparison for several throttle senings at Mach 0.33.

idicated that there were no effect at hig subsonic Mach numbers for the deflected fin. However, data presented in Figure 9 indicate that the fins are more effective at low Mach numbers and high throttle settings. This indicates that engine exhaust plume increases

baseline model consisted of a 7 inch airframe with flow-through inlets and exhaust and was externally identical with the exception of the bulged engine section and the larger wiring

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W Table 2: Longitudinal stability summary.

I I II I I I Per degree of Pitch deflection. 8. I I I - Cfg. 6 Cm,* Cn,* C%* Cma Cn, CdCn

C w o Cn,o cm, 0.00 0.2762 -0.270 0.137 .._ ... ... -0.439 0.9096 -0.454

I -22.60 8.4139 -1.766 8.092 -0.3794 0.0860 -0.3633 -0.122 0.8265 -0.175

NORMAL FORCE COEFFICIENT, CN

Figure 8. Longitudinal stability comparison for several throttle settings at Mach 0.60.

Table 3: Baseline drag (in pounds) at Mach 0.318 (q=150.0 psf).

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NORMAL FORCE COEFFICIENT. CN

Figure 9. Longitudinal stability comparison for several throttle settings with a 22.6O pitch I command at Mach 0.33.

fin loading at low Mach numbers. Thi - phenomena could provide the needed control authority for a vertical soft launch (no solid propulsion booster) when higher thrust engines of this size become available.

Enyine Performance The baseline flow-through model was

equipped with internal flow screens for simulating different engine throttle settings. Analysis of axial data obtained on the flow- through baseline configuration indicates that the drag differences between the 0%, or un-blocked,

Table 4 Body Skin I Flow Temperatures for the William Engine at 42 kRPM.

Table 5 Body Skin I Flow Temperatures for the Williams Engine at 52 kRPM.

Table 6: Body Skin I Flow Temperatures for the Sundstrand Engine at 95 kRPM.

and the 50% blocked case is on the order of 0.80 pounds with a 0.13 pound uncertainty. The zero angle of attack data are presented in Table 3, and were acquired at several roll positions, at a constant dynamic pressure of 150.0 psf. No data were acquired for the 100% blocked inlet case.

Uncorrected drag is presented in Figures 10 through 12 for various throttle settings at Mach 0.33, 0.50 and 0.60, and the net thrust coefficient is presented in Figures 13 through 15. Since there was undetermined amount of flow through the turbojet engine, there is some uncertainty associated the engine-off drag data.

Additional engine performance data,

TOT ANQLE OF ATTACK.ALPHAT, OEQ Figure 10. Drag coefficient for several throttle settings at Mach 0.33.

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om

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9 -

? m ,a* I O A * * m I* LW L(* t w no(

TOT ANOLE OF ATlACK.ALPHAT, DEQ

Figure 11. Drag coefficient for several throttle settings at Mach 0.50

~TANQLEOFA~ACK,ALPHAT, DEG

Figure 12. Drag coefficient for several throttle settings at Mach 0.60.

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E g , m

g an0

z om

0 1-

li

c Y

2 &=

*m -nu .7m a* Y O IP '* 760 (101

TOT ANQLE OF ATTACKALPHAT. DEQ

Figure 13. Thrust coefficient for several throttle settings at Mach 0.33.

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9.w

om

'Yo. .,.%I +.w .*,%I .* 2 0 1 1 1.11 3o.m

TOT ANGLE OF ATTACK.AIPHAT, OEG

Figure 14. Thrust coefficient for several throttle settinos at Mach 0 ~ 5 0

TOTANQLEOF ATTACKALPHAT. DEQ

Figure 15. Thrust coefficient for several throttle settings at Mach 0.60.

including installed fuel consumption, and comparisons with the engine cycle deck are presented by Lilleyo).

Afterbodv Heating Tables 4 through 6 present typical

afterbody thermocouple data at zero angle of attack and indicate aftbody heating decreases with increasing Mach number. This observation is reinforced by static data collected during model check-out at MICOM. This indicates that mixing between the exhaust plume and freestream tend to cool the afterbody and no plume re-attachment occurred. Additional data and analysis is presented by ~ i l l e y ( ~ ) .

CONCLUSIONS The objectives of the test were to

determine if problems exist in stability and control, in engine performance, or in heating due to aerodynamic interactions with the plume. These objectives were satisfactorily met during testing. Based on the data obtained, the following conclusions were drawn:

Basic longitudinal stability data were compared with previous wind tunnel test data of similar configurations indicating that the engine has no adverse effect on missile stability or control authority.

The engine exhaust plume increases fin loading and serves to improve missile control authority at very low Mach numbers. This effect could be very beneficial to future soft launch turbojet powered missiles.

The worst case in terms of aft heating occurs at high throttle settings and low freestream (or static) velocities. The data obtained during this investigation indicate that there is significant mixing between engine exhaust and the freestream air. However, a significant portion of the afterbody is exposed to elevated temperatures (> 200' F), and at high angles of attack (> 14') local heating can approach 400' F.

Future flow-through model investigations

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and engine-on investigations should make provisions to obtain data for the full-blocked inlet.

Future afterbody heating studies can be addressed by exhausting the engine over the afterbody in a static environment.

REFERENCES 1. Lilley, J.S., Pengelly, S.L., Fisher, P.,

“The PC as a Generic Fuel Control for Expendable Turbojet Engines”, AIAA Paper 91-3399, June 1991.

2. Convert, E.E., Thrust and Drag Its -, AIAA Volume 98.

3. LiIley, J.S., Pengelly, S.L., “Wind Tunnel Evaluation of Mid-airframe Mounted Tactical Missile Turbojet Sustainer”, AIAA Paper 92-3752, July 1992.

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