A Base Drag Reduction Experiment on the X-33 Linear Aerospike SR-71 Experiment (LASRE) Flight Program

NASA/TM- 1999-206575

A Base Drag Reduction Experiment

on the X-33 Linear Aerospike SR-71

Experiment (LASRE) Flight Program

Stephen A. Whitmore and Timothy R. Moes

Dryden Flight Research Center

Edwards, California

March 1999

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NASA/TM- 1999-206575

A Base Drag Reduction Experiment

on the X-33 Linear Aerospike SR-71

Experiment (LASRE) Flight Program

Stephen A. Whitmore and Timothy R. Moes

Dryden Flight Research Center

Edwards, California

National Aeronautics and

Space Administration

Dryden Flight Research Center

Edwards, California 93523-0273

March 1999

Acknowledgments

The authors thank the SR-71 crew for allowing access to the aircraft during a high-pressure time near the end of the program. The authors also acknowledge the expert assistance of Dale Hilliard and Jerry S. Reedy of Kaye and Associates in applying the gritted paint to the flight experiment.

NOTICE

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(301) 621-0390 (703) 487-4650

A BASE DRAG REDUCTION EXPERIMENT ON THE X-33 LINEAR

AEROSPIKE SR-71 EXPERIMENT (LASRE) FLIGHT PROGRAM

Stephen A. Whltmore, Timothy R. Moes t NASA Dryden Flight Research Center Edwards, California

Abstract area of gap between reflection plane and

Agap model, ft 2 Drag reduction tests were conducted on the LASRE/ A wetted area of forebody surface grit, ft 2 X-33 flight experiment. The LASRE experiment is a grit flight test of a roughly 20-percent scale model of an linear acceleration vector, measured at Arneas X-33 forebody with a single aerospike engine at the rear.

instrument package, ft/sec 2 The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for projected area of engine ramp onto y-z Aramp reducing base drag by adding surface roughness along plane, ft 2 the forebody. Calculations show a potential for base drag reductions of 8 to 14 percent. Flight results LASRE forebody wetted area, ft 2 n_,et corroborate the base drag reduction, with actual reference span reductions of 15 percent in the high-subsonic flight Bref regime. An unexpected result of this experiment is that base drag coefficient, referenced to base CDbase drag benefits were shown to persist well into the area supersonic flight regime. Flight results show no overall net drag reduction. Applied surface roughness causes predicted base drag coefficient, CDba_e[Moo] forebody pressures to rise and offset base drag referenced to base area reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the predicted base drag coefficient, Dbase LASRE drag experiments are inconclusive and more incompressible flow conditions, work is needed. Clearly, however, the forebody grit referenced to base area application works as a viable drag reduction tool.

forebody pressure drag coefficient, CDfore

Nomenclature

referenced to base area cOiSc) total base area for LASRE model, ft 2 total viscous forebody drag coefficient, Abase Dfore referenced to base area projected area of LASRE boat tail base Aboat onto y-z plane, ft 2 total pressure drag coefficient for the CDp LASRE model, referenced to base area A projected area of engine plug base onto eng base y-z plane, ft 2 LASRE parasite drag coefficient, C O parabase referenced to base area projected area of engine fence onto y-z Afence zero-lift drag coefficient of the LASRE plane, ft 2 CD o model, from balance, referenced to base area *Vehicle Aerodynamics Group Leader, Senior Member, AIAA. N predicted zero-lift drag coefficient of the ?Aerospace Engineer, Member, AIAA.

CD o LASRE model, referenced to base area Copyright © 1999 by the American Institute of Aeronantics and Astronautics, Inc. No copyright is asserted in the United States _mder zero-lift drag coefficient of the LASRE

Title 17, U.S. Code. The U.S. Govermnent has a royalty free license CJoP)

to exercise all rights under the copyright clahned helein for Govern model, from pressures, referenced to mental pro'poses. All other rights are reserved by the copyright owner. base area American Institute of Aeronautics and Astronautics forebody skin friction drag coefficient, psid differential pressure, lb/in 2 C f base referenced to base area freestream static pressure, lb/ft 2 P_ (rough) skin friction coefficient for rough flat CfL weighting function for surface pressure plate, referenced to Aw, et qi measurement (sin) skin friction coefficient for smooth flat CfL plate, referenced to Aw, et Reynold's number based on length Re L C pressure coefficient offset from SR-71 instrument package to Rmodel P model center of gravity, ft integrated surface pressure coefficient "_P samples-per-second sps integrated engine base pressure "_Pbase coefficient planform reference area Sref integrated boat tail pressure coefficient "_P boat reflection exit velocity, at base of model, Vbase ft/sec integrated lower engine fence pressure "_P fence coefficient Voo freestream velocity, ft/sec integrated forebody pressure coefficient "_P fore x longitudinal coordinate, fl, in.

pressure coefficient measured at i'th Cp i lateral coordinate, ft Y pressure port Z vertical coordinate, ft integrated left-nozzle ramp pressure "_ left coefficient AC Oisc) increment in total viscous forebody drag Dfore coefficient caused by added forebody integrated right-nozzle ramp pressure "_ right coefficient roughness, referenced to base area true force vector acting on LASRE base drag reduction caused by added Faero AC Dbase model, lbf forebody roughness, referenced to base area friction force acting between reflection FfRa,,, plane and model, lbf equivalent sand-grain roughness of K S surface extrusions, in.

raw force vector measured by LASRE Fra_, model balance, lbf weighting function scale factor v i i port index local flow density, at reflection plane exit Pbase at base of model, slug/ft 3 L length, ft freestream flow density, slug/ft 3 Poo mass of the LASRE model, excluding remodel reflection plane, slugs vehicle angular velocity vector, rad/sec divergence drag rise Mach number Mdi_ vehicle angular acceleration vector, rad/sec 2 freestream Mach number M slope of model surface along x-y N number of ports used in integration direction at i'th port base pressure, lb/ft 2 Pbase slope of model surface along x-z absolute pressure, lb/in 2 psia direction at i'th port Atnelican Institute of Aeronautics and Astronautics Introduction ..............................................................................................................................

Current proposed shapes for reusable single-stage-to- orbit vehicles like the Lockheed Martin X-33 and VentureStar TM reusable latmch vehicle have extremely large base areas when compared to previous hypersonic vehicle designs. 1 The comparatively large base areas for the X-33 and VentureStar TM are a consequence of the lifting-body shape of the vehicle, and the need to fit the rectangular linear aerospike engines into the base region. As a result, base drag---especially in the transonic flight regime--is expected to be quite large.

Alternatively, the need for a low-drag profile for the ascent phase of the flight has resulted in a relatively clean, low-camber forebody shape for the X-33.

Consequently, at low angles of attack one would expect the forebody drag of the X-33 to be relatively low; and that base drag would dominate the vehicle drag characteristics.

The unique configuration of the X-33, with its large base area and relatively low forebody drag, offers the Figure 1. The LASRE pod mounted on top of the SR-71 potential for a high payoff in base drag reduction. This aircraft.

paper presents results of a base drag-reduction test, conducted on the X-33 Linear Aerospike SR-71 Experiment (LASRE).2 This flight experiment attempted testing of the LASRE without having actually fired the to reduce base drag by increasing forebody surface rocket engine in flight.

roughness. This report presents results of the experiment, and compares the resulting low angle-of- The model is mounted onto the aircraft so that the attack drag numbers to the X-33 wind tunnel data base.

lateral axis is aligned parallel to the normal axis of the Effects of the aerospike rocket firing on the base drag SR-71. This alignment causes the angle of sideslip for characteristics are not addressed.

the SR-71 aircraft to be equivalent to angle of attack for Use of trade names or names of manufacturers in this the LASRE model. Thus, with a zero-angle-of-sideslip document does not constitute an official endorsement of flight condition for the SR-71 aircraft, the model is such products or manufacturers, either expressed or essentially flying at zero angle of attack. To achieve implied, by the National Aeronautics and Space better flow quality, a reflection plane was mounted Administration.

between the SR-71 and the model. The reflection plane shields the model from the SR-71 flow field.

Background on the LASRE Flight Experiment Model mold lines are constructed from a 30-in.

The LASRE experiment is a flight test of a roughly diameter cylinder which is swept away from the 20-percent half-span scale model of an X-33 forebody longitudinal axis by an angle of 20 °. At the nosetip, the with a single aerospike rocket engine at the rear. As cylinder is faired smoothly with a 15-in. radius shown in figure 1, the entire test model is mounted on hemisphere. Figure 2 shows a three-view line drawing top of an SR-71 aircraft. It was intended that LASRE of the model and documents the primary geometrical flight test data would be used to define the aerospike components--the forebody, boat tall, nozzle ramps, engine performance under realistic flight conditions and base plug, and engine fences. Figure 3 compares outer to determine plume interactions with the base and mold-lines of the LASRE to a 20-percent scale top-view engine cowl areas. NASA Dryden recently concluded of the X-33. Comparisons show a fairly close match.

Table 1 compares some vital geometric properties of the LASRE model to those of the X-33.

TMVentuleStar is a legisteled tradelnark of Lockheed Martin, Inc., Mountain View, California.

American Institute of Aeronautics and Astronautics Table 1. Comparison of the LASRE and X-33 reference dimensions. 2 Rear view Front view B at tail Symbol Description X-33 LASRE TPS

-f-

Right engine _57.8 in. _-_ Sre f Planform reference 1608 ft 2 32.15 ft 2 area Lre f Reference length 63.2 ft 13.12 ft L:oftz_/g irne p nozzle ramp Br4. Reference span 36.6 ft 3.75 ft Y_ (60 percent of Lr4. ) 30 in. _-_ 10.25 in. -_ _ X-Engine n°zzlebase plug A_.et Wetted area 5120 ft 2 101.62 ft 2 (excluding base) Abase Base area 466.9ft 2 12.04 ft 2 Note: LASRE reference data are for a half span vehicle.

Instrumentation and Processing of

., 140in. "l / I

the Onboard Measurements

I-_ 165 in. --/_t / In order to measure performance of the Linear Engine nozzle fence ---_ Aerospike engine under a variety of flight conditions, y Top view the model was motmted to the SR-71 with a pylon that F°reb°dy 7 ,_._ was instrumented with 8 load cells oriented to allow a I_/ :::,'_ i_" / /i /_ I / %1 six-degree-of-freedom measurement of the total forces

£_ / (= x / / / I/ \

and moments. The model was also instrumented with .... =-.-...._.: ........ 2>- .......... "........... _.............. I J I surface pressure ports on the forebody, boat tail, base, 980551 engine ramps, and the lower engine fence.

Figure 2. The LASRE test model.

Other onboard instrumentation included the alrdata measurements--Mach number, airspeed, angle of attack, angle of sideslip, and altitude--from the onboard alrdata system of the SR-71, and vehicle accelerations and angular rates from strapdown sensors located near the vehicle center of gravity. All onboard analog instrumentation were sampled using 12-bit pulse code modulation (PCM) and telemetered to the ground for postflight analysis. The airdata parameters were LASRE mold telemetered and recorded at 50 samples-per-second (sps). Onboard accelerometer and rate-gyroscope readings were telemetered and recorded at 200 sps.

Force Balance Data Measurements X-33 mold lines 980552 The force balance measurements consisted of 8 load cells, oriented to give outputs proportional to the forces Figure 3. A comparison of the LASRE outer mold lines acting along the axial, vertical, and lateral directions on with the X-33.

the balance (fig. 4). A calibration tensor measured by Lockheed Martin (Palmdale, California) prior to delivering the LASRE experiment to NASA Dryden Amelican Institute of Aeronautics and Astronautics reflection plane), Faero is the vector of corrected -Top mounting flange aerodynamic loads acting on the model, Fra w, is the force vector calculated from the tmcorrected load data, Ameas is the measured linear acceleration vector, co is Load cell the angular rate of the vehicle, 63 is the angular balance acceleration of the vehicle, and Rmocle I is the vector distance from the location of the SR-71 aircraft Axial accelerometer package to the center of gravity of the Lateral load model. The center of gravity of the LASRE model lies 39.025 ft aft, 7.408 ft above, and 2.708 ft inboard of the SR-71 accelerometer package. For the SR-71 LASRE experiment, angular acceleration was not directly measured; instead angular acceleration was computed by numerically differentiating the angular rate vector. 4 Surface Pressure Measurements LBottom ig pylon mounting flange (to SR-71) Pressure instrumentation consisted of flush pressure (to model) 980553 taps distributed on the forebody, boat tall, engine ramps, engine base plug, thruster cowling, and engine fences. A Figure 4. Schematic of the LASRE force balance.

total of 95 ports were distributed on the forebody and boat tail. Locations of the forebody and boat tail ports are shown in figure 5. In addition 58 ports were located was used to relate the output readings to the true forces in the engine base area, with 20 pressure ports located and moments acting on the balance. The balance was on the left engine ramp, 22 ports on the right engine not re-calibrated during the course of this flight ramp, and 16 ports on the engine base plug. An program.

additional 2 pressure ports were located on the trailing edge of the lower engine fence. Figure 6 shows the Raw force-balance data were sampled at 50 sps, and locations of the engine pressure ports.

these were low-pass filtered using a second-order Butterworth digital filter 3 to remove noise caused by Forebody, boat-tail, and nozzle surface pressures were structural vibrations and aerodynamic turbulence. Filter sensed using electronically scanned pressure (ESP) latency was accounted for by time-skewing the data modules. Because of pressure ranges expected during after filtering. The filtered data were corrected for zero- aerospike engine hot-fire tests, engine ramp and fence offsets using preflight and postflight zero-tare data. The pressures were sensed using _+50 psid pressure sensors; zero-readings were taken for each load cell by averaging all other surface pressure measurements were made one minute of data each, from both preflight and using _+10 ESPs. All ESPs were referenced to a highly postflight. The calibration tensor was then used to accurate 0-38 psia 20-bit digital pressure transducer.

compute the axial, normal, and side loads, and pitch, The reference pressure was added to the differential ESP roll, and yaw moments acting at the balance.

readings to determine the absolute local pressure reading. Temperature environments of the ESP were To determine the true aerodynamic forces acting on controlled using heater blankets. Zero-shift corrections the model, it is necessary to remove the centrifugal force using preflight and postflight tare readings were also and vehicle accelerations acting at the model center of performed. To reduce the effects of structural vibrations gravity. These corrections were computed using the and aerodynamic turbulence, pressure measurements strapdown instruments onboard the SR-71 aircraft. The were digitally filtered. All pressure data were measured vector equations for the force transformations are at 50 sps.

Faero = Fra w,-{mmode I"

Flight Test Maneuvers

(1) [Ameas + [60 × 60 × Rmode I + 63 × Rmodel] ] } Acceleration data from subsonic to supersonic flight conditions were used in this analysis. Initially, level In equation 1, mmode I is the mass of the model, (the part of the total experiment mounted above the altitude accelerations were flown for envelope Amelican Institute of Aeronautics and Astronautics 8O i Front view,.

Upper engine fence 30 in.

i looking aft _ 6O

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10.25 in.

Z, Lowerengine fence 980555 in.

2o.............. i.......... i......

Figure 6. Layout of LASRE engine nozzle plug and o ...... ramp pressure port.

2O Baseline Drag Measurements on the Top viewl looking down LASRE Model Configuration Baseline drag measurements on the clean LASRE y, 0t in.

configuration will be presented first. The clean configuration is defined as the model without added - 20 forebody surface roughness. Data derived from four - 40 typical flight maneuvers performed during flights 46, 0 20 40 60 80 100 120 140 160 47, 48, and 49 are used to illustrate the drag properties x, in.

980554 of the model. These baseline drag data verify the Figure 5. Port locations on LASRE forebody and resolution, repeatability, and accuracy of the boat tail.

measurements; and substantiate the earlier assertions that base drag is the dominating drag-force component.

In the remainder of this paper, all drag coefficient data expansion and flutter clearance. Once the flight will be referenced to the base area of the LASRE model envelope clearance was obtained, a more fuel-efficient as presented in table 1.

dipsy maneuver was used to accelerate through the large transonic drag rise. The dipsy maneuver began at Overall Model Drag Measurements 28,000 ft and Mach 0.9. The pilot put the aircraft into a slight dive to help get through the transonic drag rise Figure 7 shows the overall drag coefficient, CDo, for and then leveled the aircraft at approximately Mach 1.07 the clean LASRE model plotted as a function of Mach and an altitude of 25,000 ft, which was the minimum number. Repeatability of the data are excellent, having a altitude cleared for transonic flight. The aircraft total scatter band of less than 0.015. For comparison continued to accelerate at an altitude of 25,000 ft until it purposes wind-tunnel derived values for the X-33 total obtained an equivalent airspeed of 450 kn, at which point the pilot initiated a constant equivalent airspeed CDo are also plotted. The very large transonic drag rise climb to the desired Mach number. Structural load observed on the flight data does not show up on the restrictions on the LASRE experiment required that the wind tunnel predictions. Reasons for the transonic drag angle of sideslip--equivaient to angle of attack in the difference are not definite at this point; however, it is model axis--be restricted to less than two degrees.

possible that this difference is an effect of the sting- Because of this restriction, all of the drag data obtained mount used to support the X-33 wind tunnel model.

are essentially for the zero-lift flight condition--CD0.

American Institute of Aeronautics and Astronautics pressure port distribution on the LASRE model is not LASRE total drag coefficient dense enough to allow a full three-dimensional from force balance geometric integration of the pressures. If a geometrical Mach number Flight range grid were used to numerically integrate the pressures, o 46 0.78to 1.54 the uneven port spacing would give far too much area o 47 0.70to 1.52 o 48 0.68to 1.62 weighting to the ports located in the sparsely populated 49 0.62 to 1.78 regions. Instead, for a given geometrical component --,,- X-33 total drag coefficient, (such as the forebody surface) the surface integral was wind tunnel mechanized as a weighted average of the measured *Referenced to LASRE pressures.

base area 1.0 N .8 [qiCpi] .6 (2) -_pi=l Zero-lift drag .4 [qi] coefficient,* i=1 CD0 .2 Instead of weighting pressures by their local area, the weighting function applied in equation 2 is the projection of the local surface onto the y-z plane, -.2 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 Mach number 980556 v i

(3)

Figure 7. Baseline LASRE total zero-lift drag qi = j (3X_ 2 (3X_ 2 coefficient. 1 + k byji + k bzji Equation 3 weights more heavily ports that are Individual Components of the Overall Model Drag aligned more perpendicular to the drag axis. The Coefficient numerical integration was performed for 6 geometrical components on the model: The shape of the LASRE CDo curve as a function of Mach number can be better understood by examining 1. the model forebody, aft to 140 in. behind the the individual drag-force components acting on the nosetip, model. Since the LASRE model has no camber and 2. the engine nozzle left ramp, nominally flies at zero local angle of attack, induced 3. the engine nozzle right ramp, drag-due-to-lift is considered to be negligible. Thus there are 3 remaining drag components which must be 4. the engine nozzle base plug, considered as important: 5. the LASRE model boat tall, and 1. Base and boat tall drag, 6. the lower engine fence.

2. Forebody pressure-profile drag, and In equation 3, v i is an arbitrary weighting function 3. Viscous drag from forebody skin-friction and scale factor which was assigned to give better rtm-to-run residual parasite drag.

data consistency. For the base, ramp, boat tail, and fence integrations, the value of v i was always unity. For the Effects of each component on the total LASRE drag are forebody integration, ports along the model centerline, now presented.

and on the fiat side-falrings unity values for v i were assigned. Ports along the sides of the swept cylindrical Surface Pressure Integration forebody were assigned values of v i = 1.5. This Forebody, boat tall, and nozzle base drag coefficients weighting increment helped to account for the are computed by numerically integrating the pressure sparseness of ports along the swept cylindrical sides of measurements along the surface of the body. The the forebody.

Amelican Institute of Aeronautics and Astronautics Oncethe individual pressure coefficients of each constant at approximately 0.38 until the divergence drag

geometrical component are determined, surface

rise Mach number, Mdi_, of approximately 0.90 is

pressure drag iscalculated asthearea-weighted average

reached. After the divergence Mach number is reached,

of integrated pressure coefficients for individual

compressibility effects dominate and base drag geometrical components, coefficient rises rapidly. Beyond Mach 1, base drag m drops off steadily. In the subsonic flight regime, base CDp = Cpfo,.e - drag accounts for approximately 125 percent of the overall model drag. Approximately 80 percent of the transonic drag rise can be attributed to compressibility effects on base drag.

(4)

+ Aboa,CE boo,+ Aj.e.ceG, ..... 1 /

Since base drag is higher than overall model drag for subsonic flight conditions, one would expect a substantial amount of forebody suction to occur. The [2Aramp + Aeng base + Aboat + Afence I lower curve in figure 8 verifies this expectation. The forebody drag coefficient is negative until the transonic The resulting base drag coefficient, CDbas e = --Cpbase, drag rise is encountered. Even in the transonic flight regime, forebody drag coefficient accounts for less than and forebody pressure drag coefficient, CD_,.e_ = Cplo,.e, are presented as a function of Mach number in figure 8. 8 percent of the total model drag coefficient. The strength of forebody suction is likely a result of a clean For comparison purposes a fairing of the total drag forebody shape for the LASRE. As mentioned coefficient, derived from figure 7, is also presented. In previously, the mold lines for the LASRE forebody are a the subsonic flight regime base drag remains relatively 20 ° swept cylinder faired to flat sidepanels. This shape ensures that a significant adverse pressure gradient does not occur along the forebody.

Base + forebody, This premise is illustrated in figure 9(a) where the integrated pressures forebody pressure distribution at Mach 0.70 is plotted as Mach number a function of the vertical (z) and longitudinal (x) Flight range = 46 0.78 to 1.54 coordinates. Figure 9(b) shows locations of the pressure = 47 0.70 to 1.52 ports on the forebody. From the nosetip to o 48 0.68 to 1.62 approximately 40in. aft, the pressure gradient is = 49 0.62 to 1.78 strongly favorable. Between 40 in. and 100 in. aft, the Total drag fairing, force balance pressure gradient is almost fiat; and beyond 100 in. aft, *Referenced to LASRE base the pressure gradient becomes strongly favorable again.

area 1.0 Although the surface pressure gradient between 40 in.

and 100 in. aft is approximately neutral, the boundary .8 layer in this region is clearly turbulent 5 and flow separation is very unlikely. Pressure distributions for other Mach numbers have a similar profile.

Drag .4 Skin Friction and Parasite Drag Coefficients coefficient* Total drag coefficient, CDo , is compared with overall pressure drag coefficient, CDp , in figure 10. Residuals between the two curves are also plotted. Obviously, residual data include measurement errors in both the -.2 .2 .4 .6 ,8 1,0 1,2 1,4 1,6 1.8 2.0 Mach number force balance and surface pressure data; however, the 980557 residual data represent a crude measure of the combined Figure 8. Comparison of the total LASRE drag viscous 6 drag forces acting on the model. As will be coefficient with the base and forebody pressure drag coefficients.

shown in the next section, these viscous forebody forces Amelican Institute of Aeronautics and Astronautics Base + forebody, 0 C ) -> z = 1.50 in.

integrated pressures [] C ) -> z = 2.90 in.

Mach number ) -> z = 9.20 in.

Oc

Flight range A C ) -> z = 12.70 in.

o 46 0.78 to 1.54 t_C ) -> z = 15.60 in.

D 47 0.70 to 1.52 dC ) -> z = 21.90 in. o 48 0.68 to 1.62 A 49 0.62 to 1.78 u'c ) -> z = 35.00 in.

Total drag fairing, force balance _c ) -> z = 46.00 in.

........ Computed viscous forebody Cp -> z = top row drag coefficient *Referenced to LASRE 2.0 I Flight 46, M_ = 0.70 base area

I

**Parasite residual, total drag, integrated pressures 1,0 1.5 [.....................................................................................................

1,0. "..................................................................................................

.8 Forebody pressure .5 ...................................................................................................

.6 coefficient Drag .4 coefficient*

.2

-.5 -1.0 50 100 150 -.2 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 Distance aft, x, in.

980558 Mach number 980560 (a) Forebody pressure distribution.

Figure 10. Comparison of the total LASRE drag coefficient with total pressure drag coefficient.

See legend on Fig. 9(a) for z-axis measurements indicated by connected pressure points gap between the lower side of the model and the 4O reflection plane. The ram drag is considered as equivalent to the parasite drag which forms on more .z, 20 In.

complex aircraft configurations.

The forebody skin friction coefficient (referenced to the base area of the LASRE model) was evaluated by 0 20 40 60 80 100 120 140 160 numerically solving the nonlinear equation for the x, in.

980559 Schoenherr line, 7 (b) Pressure ports on side view of forebody.

Figure 9. LASRE forebody pressure data, Flight 046, Mach 0.70.

IC fba_ _ Abase

(5)

Abase strongly influence the base drag. As a check on the accuracy of this crude viscous drag measurement, an where, Re L is the forebody Reynold's number, Abase is estimate of the viscous forebody drag coefficient, the base area, and Awe t is the wetted area of the c(_isc) forebody (table 1).

C_io,.eisc), is also calculated. For the LASRE model Dio,.e has two principal components: (1) the forebody skin The parasite drag (referenced to the base area of the friction drag and (2) the ram drag resulting from a 1-inch LASRE model is calculated by performing a Amelican Institute of Aeronautics and Astronautics

one-dimensional momentum andforcebalance in the

Integrated pressures + viscous

axial direction

drag estimate Mach number Flight range V 2 = o 46 0.78to 1.54 Agap(P_ V2- Pbase base) (6) o 47 0.70to 1.52 o 48 0.68to 1.62 2F fRam- Agap(P_ - Pbase) A 49 0.62to 1.78 Total drag fairing, force balance In equation 6, Agap is the frontal projection area of the *Referenced to LASRE base area gap between the model and reflection plane, and FfRam is the skin friction force acting between the reflection 1.0 _) T_tal diag coeffi_ien i i plane and the lower surface of the model. Normalizing by freestream dynamic pressure and LASRE base area, equation 6 becomes co$flriaigent , .6 .4 : : 0-+-----i-----i ..........

V 2 V 2 (P_ _-Pbase base) 1 V 2

72P_ ,2

(7) 1 _Residual • i :: :: V FfR_m Abase (P_- Pbase) Drag 01 ........... i........... i .......... I 1 V 2 A 1 V 2

coefficient* -.11 ........... ibiEJrage0_ff[ei_nt m sid _ai .......... I

_P_ _ base Agap _P_ -.2' .... ' .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 Mach number 980561 Assuming that exit velocity is much smaller than entrance V 2 V 2 Figure 11. Comparison of total surface pressure velocity, P_ _ >> Phase base, and defining 2 coefficients.

CDp_,.%_e =- F ft_am/(Ab as e P_ V /2) equation 7 reduces to estimates is approximately 0.015, and the maximum - Agap F1 1C

(8)

deviation is 0.04. Because there are more uncertainties

CDparabas e _k "4"_ Dbase]

involved in deriving the estimate of CD(P), it is likely that As mentioned earlier, total viscous forebody drag pressure-derived total drag coefficient estimates coefficient is the sum of the skin friction and parasite contribute a larger portion of the overall error-- drag coefficients (referenced to base area) especially in the transonic flight regime.

cOiSc)

Development of a Drag Reduction Strategy

D for e Cfbas e + (9) = CDparabas e The data presented in figures 8 through 11 clearly cOiSc) Dfor e is also plotted on figure 10. The computed support earlier assertions that base drag dominates the values show reasonable agreement when compared to overall drag LASRE. For subsonic conditions Saltzman 1 the residual data.

and Hoerner 7 have demonstrated a well-defined Comparison of the Drag Coefficients Computed Using C Oisc) for vehicles correlation between D for e and C Dbas e the Two Methods with a wide variety of shapes, sizes, and base-to-wetted cOiSc) area ratios. For two-dimensional shapes Hoerner 7 has If calculated values for Dfor e are added to integrated demonstrated that the subsonic correlation is pressure drag, C D , an estimate of total model drag approximated by the empirical formula coefficient, CD(P),Pis-- generated independently of the force balance measurements. The two independent drag .135 coefficient estimates are compared in figure 11.

CDbas e -- 3 C/-_ isc) (10) Residuals between the two estimates, CDo - CD(P), are _ D fore also plotted. The average difference between the two American Institute of Aeronautics and Astronautics For three-dimensional shapes, the correlation formula is 0.029 CDbas e - C[__._isc ) (11) I_ _ D fo,. e Saltzman 1 has found that for large-scale reentry-class flight vehicles the two-dimensional equation is a more accurate representation of the flight data. Based on this reasoning, equation 10 will be preferred in this analysis.

The reasons for the correlation predicted by equations 10 and 11 become more clear if one examines flow visualizations images of the LASRE obtained in the NASA Dryden Flow-Visualization Facility. 8 Figure 12 shows water-tunnel flow images taken from tests of a 2.5-percent scale model of the LASRE/SR-71 configuration. Although the Reynolds numbers for the water tunnel tests (N1000) are significantly lower than for flight (N2-5 x 106), nevertheless, the images presented serve as a good illustration of the LASRE base flow characteristics in the absence of engine thrust. The images clearly show the external freestream flow pumping fluid away from the engine base. This pumping effect reduces base pressures significantly. The forebody boundary layer arriving at the edge of the model acts as an insulating layer between the external flow and the separated base area. This insulating layer reduces the effectiveness of the pumping mechanism. Because the thickness of the forebody boundary layer is directly 980562 related to the viscous forebody forces, the source of the correlation of equation 10 becomes evident.

(a) Top view.

980563 (b) Right side view.

Figure 12. Water tunnel flow visualization images for a 2.5-percent scale LASRE model.

Amelican Institute of Aeronautics and Astronautics The above discussion leads to a possible method for -- Base drag, force balance base drag reduction by increasing the viscous drag flights 46 to 49 ...... Computed base drag, Hoerner acting on the forebody of the vehicle. This viscous drag correlation model, baseline increase serves to increase boundary thickness and *Referenced to LASRE reduces the effectiveness of the vacuum-pump acting at base area 1.0 the base. If the boundary layer modification can be Base _drag performed without additional flow separation or excessive streamline displacement along the forebody, it .8 may be possible in some instances to decrease the drag of the entire configuration.

Drag .6 Development of a Mathematical Model for the LASRE coefficient* Drag Coefficient .4 To determine whether this concept is feasible or not, a mathematical model of the LASRE base drag coefficient cOiSc) must first be developed which has Dfor e as a .2 parameter and accounts for flow compressibility. As CDfore mentioned earlier, LASRE base drag data show that in increase, percent constant until the divergence Mach ntLmber of Predicted base drag reduction .10 approximately 0.90 is reached. After this point compressibility effects dominate and base drag coefficient rises rapidly. Beyond Mach 1, base drag ACDbase .05 ......... _ .............

drops steadily. These trends suggest a base drag

'-_----. .... "-'7. _._--. --_

compressibility function of the form 0.6 .8 1.0 1.2 1.4 1.6 1.8 Mach number 980564 ~ 7_ (o) _ .135 Figure 13. Comparison of the LASRE base drag M < Mdi _ =:> COba_e[M ] = t_Oba_e 3c_isc ) coefficient with base drag prediction.

_ Dfore N Mdi _<M <1=:> CDb_[M ] = cOiSc) M2i_ 1 7.(o) F on the model of equation 12) if Dfo,.e is increased by 25 percent, 50 percent, 75 percent, and 100 percent Dbase respectively.

1 =(o) [-

+ (12)

The mathematical model of equation 12 can be

6Dbo [M ] = extended to total drag coefficient by adding in the

viscous and forebody pressure-drag terms 2 1_ (o) F ~ __ cOiSc) + CDo = Cpfo,.e + Dfo,.e CDbase[Moo] (13) The analytical drag model of equation 13 is compared The elements of equation 12 are derived from equation with the measured LASRE base drag data in figure 14.

10 with modifications for compressibility defined by the Again, for such a simple model the comparison shows Karman-Tsien correction, 9 and rules of similarity for good agreement.

transonic flow. 1° The base drag model of equation 12 is Increasing the Forebody Viscous Drag by Increasing compared against measured LASRE base drag data in Surface Roughness figure 13. For such a simple model the agreement is reasonable. Also presented in figure 13 are base drag Clearly, one of the most convenient methods of reduction increments that would be expected (based increasing the forebody viscous drag is to add roughness American Institute of Aeronautics and Astronautics -- Total drag fairing, force balance, flights 46 to 49 .... Computed drag, Hoerner correlation model, baseline *Referenced to LASRE base area 1.0 .8 ....................................... port .................................. ii .................... ii ..................

.6 ......................................

CD 0 .2 Figure 15. Close-up of LASRE grit application.

.6 .8 1.0 1.2 1.4 1.6 1.8 Mach number 980565 Figure 14. Comparison of the total LASRE drag Gritted surface area ~ 32.4 sq. ft., 1/3 of forebody area coefficient with total drag prediction.

to the surface. Other methods such as using vortex generators to energize the boundary layer would probably work more effectively, but their intrusiveness into the flow precludes this method for application to the hypersonic re-entry vehicle problem. For the LASRE drag reduction experiment no. 24 Silicon Carbide (0.035 in.) grit was glued to the skin using a spray-on adhesive and the surface was sealed using a high-tensile strength white enamel paint. The resulting surface, Figure 16. LASRE forebody surface grit.

depicted in figure 15, had an equivalent sand-grain roughness that varied between approximately 0.02 in.

and 0.05 in. In an attempt to avoid inducing additional When the surface of the plate is roughened, skin friction flow separation at the boat tall or along the forebody, increases considerably. For a fully rough plate the only the fiat sides of the LASRE model were gritted. The empirical formula, grit, depicted in figure 16, covered an area of 32.4 ft 2_ approximately 1/3 of the forebody wetted area.

(rough) F F L 99 2.57 (15)

cjL = L2'635 + 0'6181°ge/=-//L%dd

Surface Roughness Calculations is a good approximation. In equation 15, K s , is the In order to predict effectiveness of the surface grit in equivalent sand-grain roughness of the surface reducing base drag, calculations of the increment Oisc) extrusions. Using equations 14 and 15, the increment in in C D were performed using the method of Mills viscous forebody drag caused by added roughness is and H_aIlg. 11' 12 For a smooth flat plate of length L, the averaged skin friction coefficient is related to Reynolds (16) number according to the empirical formula AcOisc) Fc(r°ugh)_m)]Agrit Of°"e = I'fL -- C Abas----_e In equation 16, A is the wetted area of the surface c_m)= 0.0740 (14) grtt [ReL ]1/5 grit, and L is the length of the gritted area measured at Amelican Institute of Aeronautics and Astronautics thecentroid. Based onanestimated range of surface supersonic flight regime. Because base drag of

roughness from 0.02in.to 0.05 in., the calculated

supersonic projectiles had never been previously

increase in C_ isc) ranges from 18 percent to 30 percent

cOiSc) lJfore correlated to D:o,.e, the supersonic base drag over the range of Mach and Reynolds numbers reduction was a significant positive result. Figure 18(b) encountered during the LASRE flights.

shows the measured base drag reduction compared to the base drag reduction predicted using the analytical

Flight Test Results for the

Forebody Grit Experiment model (equations 12, 14, 15, and 16) assuming K s =

{0.02in., 0.05in., and 0.10in.}. Measured drag Unfortunately, the drag reduction experiment reduction shows excellent agreement with ranges occurred so late in the LASRE program that only one predicted by the analytical model.

flight test was conducted prior to the cancellation of the program. As a result, it was not possible to verify the flight-to-flight repeatability of the experiment. Figure 17 Base drag coefficients summarizes the flight results. The grit application did -- Flights 46 to 49, without grit not reduce the total drag of the configuration.

Flight 51, with grit Nonetheless, because the base drag was reduced, results *Coefficients referenced to LASRE base area of the experiment are encouraging.

.8 Flight .7 46 to 49 CD0 fairing, without grit .6 CDbase* 51 balance CD0, with grit .5 46 to 49 CDbas e fairing, without grit .4 51 CDbase, with grit .3 .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 1.4 1.5 46 to 49 CDfor e fairing, without grit Mach number 980569 51 CDfore, with grit (a) Base drag.

Skin friction increment, due to grit 1.0 .8 Base drag reduction increment Predicted, K_ s = 0.02 in.

.6 .... Predicted, K_ s = 0.05 in.

Predicted, K_ s = 0.10 in.

Drag coefficient .4 Measured flight 51 with grit

.06 _ i _ _ _ i _ i i

.2 .04.05 .......... i.......... ii 2 ".......... iiiiiiiiiill ;i;;;iFii_/;t dai; iiiiiiiiiiiill ACDbase .03 .02 .01 -.2 .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 1.4 1.5 .5 .6 .7 .8 .9 1.0 1.1 1.2 1.3 1.4 1.5 Mach number 980568 Mach number 980570 Figure 17. Effect of LASRE forebody grit: summary of (b) Drag reduction increment.

drag components.

Figure 18. Effect of forebody grit on LASRE base drag.

Base drag data are shown in greater detail in Overall drag of the configuration was not reduced figure 18. Figure 18(a) shows the base drag coefficient because the forebody grit modifications caused the plotted as a function of Mach number. Forebody grit forebody pressures to rise. The forebody pressures reduces base drag by a peak of 15 percent in the high- along the top and cylindrical sides of the model with grit subsonic flight regime. Furthermore, drag reduction and without grit are compared in figure 19(a). The port locations for the pressures being compared are shown benefits persist beyond Mach 1._-well into the Amelican Institute of Aeronautics and Astronautics in figure19(b).These pressure data,obtained from Theflightresults suggest thatthetotaldrag model of flight46(nogrit)andflight51(withgrit)atMach 0.7, equation 13must bechanged toinclude apossibility of are plotted asafunction ofthelongitudinal distance aft increasing forebodypressure drag with surface of the nosetip. Noticethat although the pressure roughness modifications. It islikelythatthe relationship distribution alongthemodel centerline wasbasically offorebody pressure drag toviscous forebody drag will unchanged, thepressures onthesides of theforebody beconfiguration dependent. Clearly, moreworkneeds aregenerally higher forthegrit-on data. Thisforebody to beperformed before moredefinite conclusions can pressure riseis further demonstrated by comparing the reached. It isalso clear, however, that forconfigurations integrated forebodypressure drag coefficients in where base drag isadominating factor, theforebody grit figure 17. When combined withadded skin-drag caused method isapotentially useful drag reduction tool.

by thesurface roughness, theforebody pressure rise

offsets thebenefits gained bythebase drag reduction. Summary_ and Concluding Remarks

A drag reduction experiment was conducted on the X-33 Linear Aerospike SR-71 Experiment.

Flight 46, Flight 51, The flight experiment performed baseline drag without grit with grit o Top row • Top row measurements on a clean experiment configuration, and o Left side ports • Left side ports then attempted to reduce the base drag by increasing the A Right side ports • Right side ports forebody skin friction using added surface roughness.

Preflight calculations showed that proposed surface roughness modifications would result in base drag reductions of 8 to 14 percent.

Moo ~ 0.70 Flight results verified the effectiveness of the surface roughness technique for reducing base drag. The peak base drag reduction was approximately 15 percent. The Forebody base drag reduction also persisted well into the pressure coefficient supersonic flight regime. Since base drag of supersonic projectiles had never been previously correlated to viscous forebody drag, the sizable supersonic base drag reduction was a significant positive result.

Unfortunately, flight test results for the rough-surface configuration did not demonstrate an overall net drag -1 reduction. The surface grit caused a rise in forebody 0 50 100 150 pressures. Coupled with increased forebody skin-drag, Distance aft, x, in.

980571 the forebody pressure rise offset benefits that were (a) Forebody pressure distribution.

gained by base drag reduction. Because the flight tests did not demonstrate an overall net drag reduction, results of the drag reduction experiment are • Top row inconclusive. It is clear; however, that with some e Side ports 6O refinement, the forebody grit method provides a potentially useful drag reduction tool.

4O

References

z, 20 in.

1Saitzman, Edwin J., Charles K. Wang, and Kenneth W. Iliff, Flight-Determined Subsonic Lift and Drag Characteristics of Seven Lifting-Body and Wing-Body 0 20 40 60 80 100 120 140 160 Reentry Vehicle Configurations With Truncated Bases, x, in.

AIAA Paper 99-0383, January 1999.

980572 (b) Forebody pressure ports, side view.

2Corda, Stephen, David P. Lux, Edward T. Schneider, and Robert R. Meyer, Jr., "Blackbird Puts LASRE to the Figure 19. Comparison of the forebody pressure distributions with and without grit. Test," Aerospace America, February 1998, pp. 25-29.

American Institute of Aeronautics and Astronautics 3Otnes, Robert K. and Lauren D. Enochsen, Digital 8Del Frate, John H., NASA Dryden Flow Visualization Time Series Analysis, John Wiley & Sons, New York, Facility, NASA TM-4631, May 1995.

1972, pp. 237-239.

9Freiberger, W. F., Ed., International Dictionary of 4Haering, Edward A., Jr. and Stephen A. Whitmore, Applied Mathematics, D. Van Nostrand Company Inc., FORTRAN Program for Analyzing Ground-Based Princeton, NJ, 1960, p. 506.

Radar Data: Usage and Derivations, Version 6.2, NASA TP-3430, August 1995.

l°Kaplan, Carl, On Similarity Rules for Transonic Flows, NACA TN-1527, Washington D.C., January 5Frieberger, W. F., ed., The International Dictionary 1948, pp. 8-10.

of Applied Mathematics, D. Van Nostrand and Company, Inc., Princeton NJ, 1960, pp. 828, 829.

11Mills, Anthony, F. and Xu Hang, On the Skin 6Hoerner, Sighard F., Fluid-Dynamic Drag, Self- Friction Coefficient for a fully Rough Flat Plate, Published Work, Library of Congress Card J. Fluids Engineering, vol. 105, September 1983, no. 64-19666, Washington, D.C., 1965, pp. 3-19, 3-20, pp. 364-365.

15-4, 16-5.

12Mills, Anthony E, Heat Transfer, Richard D. Irwin, 7Schlicting, Hermann, Boundary-Layer Theory, 7th Inc., Homewood, IL, 1992, pp. 282-328.

ed., translated by Dr. J. Kestin, McGraw-Hill Publishing Co., New York, 1979, pp. 641,642.

American Institme of Aeronautics and Astronautics

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

March 1999 Technical Memorandum

4.TITLE AND SUBTITLE 5. FUNDING NUMBERS

A Base Drag Reduction Experiment on the X-33 Linear Aerospike SR-71

Experiment (LASRE) Flight Program

WU 242-33-02-00-23-00-T 15

6. AUTHOR(S)

Stephen A. Whitmore and Timothy R. Moes

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

NASA Dryden Flight Research Center

RO. Box 273

H-2333

Edwards, California 93523-0273

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National Aeronautics and Space Administration

NASA/TM-1999-206575

Washington, DC 20546-0001

11. SUPPLEMENTARY NOTES Presented at the 37th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 11-14, 1999 as AIAA-99-0277.

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13. ABSTRACT (Maximum 200 words) Drag reduction tests were conducted on the LASRE/X-33 flight experiment. The LASRE experiment is a flight test of a roughly 20-percent scale model of an X-33 forebody with a single aerospike engine at the rear.

The experiment apparatus is mounted on top of an SR-71 aircraft. This paper suggests a method for reducing base drag by adding surface roughness along the forebody. Calculations show a potential for base drag reductions of 8 to 14 percent. Flight results corroborate the base drag reduction, with actual reductions of 15 percent in the high-subsonic flight regime. An unexpected result of this experiment is that drag benefits were shown to persist well into the supersonic flight regime. Flight results show no overall net drag reduction.

Applied surface roughness causes forebody pressures to rise and offset base drag reductions. Apparently the grit displaced streamlines outward, causing forebody compression. Results of the LASRE drag experiments are inconclusive and more work is needed. Clearly, however, the forebody grit application works as a viable drag reduction tool.

14. SUBJECTTERMS 15. NUMBER OF PAGES

Aerospike engine, Base drag, Skin friction

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A03

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