Document
NASA/TM-1998-206565
Flight Stability and Control and
Performance Results from the Linear
Aerospike SR-71 Experiment (LASRE)
Timothy R. Moes, Brent R. Cobleigh, Timothy H. Cox, Timothy R. Conners, and Kenneth W. lliff Dryden Flight Research Center Edwards, California Bruce G. Powers Analytical Services and Materials, Inc.
Edwards, California National Aeronautics and Space Administration Dryden Flight Research Center Edwards, California 93523-0273 August 1998
NOTICE
Use of trade names or names of manufacturers in this document does not constitute an official endorsement of such products or manufacturers, either expressed or implied, by the National Aeronautics and Space Administration.
Available from the following: NASA Center for AeroSpace Information (CASI) National Technical Information Service (NTIS) 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076-1320 Springfield, VA 22161-2171 (301) 621-0390 (703) 487-4650 FLIGHT STABILITY AND CONTROL AND PERFORMANCE RESULTS FROM THE LINEAR AEROSPIKE SR-71 EXPERIMENT (LASRE) Timothy R. Moes,* Brent R. Cobleigh,* Timothy H. Cox, t Timothy R. Conners,* and Kenneth W. Iliff* NASA Dryden Flight Research Center, Edwards, CA Bruce G. Powers + Analytical Services and Materials, Inc.
Edwards, CA Abstract lateral accelerometer output (positive ay toward the right), ft/sec 2 The Linear Aerospike SR-71 Experiment (LASRE) is vertical accelerometer output (positive presently being conducted to test a 20-percent-scale az down), ft/sec 2 version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single b SR-71 reference span, 56.7 ft Stage to Orbit Technology Demonstrator Vehicle. The butt line, in.
rocket engine was integrated into a lifting body B.L.
configuration and mounted to the upper surface of an c SR-71 mean aerodynamic chord, 37.7 ft SR-71 aircraft. This paper presents stability and control center of gravity, percent c results and performance results from the envelope c.g.
expansion flight tests of the LASRE configuration up to C! rolling moment coefficient Mach 1.8 and compares the results with wind tunnel rolling moment bias, coefficient estimate predictions. Longitudinal stability and elevator control CI b effectiveness were well-predicted from wind tunnel for 13 =0 ° tests. Zero-lift pitching moment was mispredicted rolling moment due to nondimensional roll transonically. Directional stability, dihedral stability, and C/p rate derivative, _ C//O(pb/2 V), rad !
rudder effectiveness were overpredicted. The SR-71 rolling moment due to nondimensional yaw handling qualities were never significantly impacted as a Ctr result of the missed predictions. Performance results rate derivative, 3 C//O (rb/2 V), rad-I confirmed the large amount of wind-tunnel-predicted rolling moment due to angle of sideslip transonic drag for the LASRE configuration. This drag CI_ derivative, OCI/O_, deg "1 increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could rolling moment due to aileron derivative, not be achieved on a hot day without depleting the CI_° _C//_Sa , deg -1 available fuel.
rolling moment due to rudder derivative, Cl_r Nomenclature OCl/_6r , deg -1 Cm pitching moment coefficient _x longitudinal accelerometer output (positive forward), ft/sec 2 pitching moment bias, linear coefficient Cm b estimate for tx = 0 ° pitching moment due to nondimensional Craq pitch rate derivative, OCrn/_(qc/2 V), *Aerospace Engineer, AIAA member.
rad 1 tAerospace Engineer.
pitching moment due to angle of attack Copyright © 1998 by the American Institute of Aeronautics and CmcL Astronautics, Inc. No copyright is asserted in the United States under derivative, 3Cm/bOC , deg -l Title 17, U.S. Code. The U.S. Government has a royalty-free license pitching moment due to elevon derivative, to exercise all rights under the copyright claimed herein for Govern- Cmae mental purposes. All other rights are reserved by the copyright owner.
_Cm/O_e , deg -1 American Institute of Aeronautics and Astronautics yawing moment coefficient acceleration of gravity, 32.174 ft/sec 2 C n g yawing moment bias, coefficient estimate /-/p pressure altitude, ft Cn b for [_ =0 ° roll moment of inertia, slug-ft 2 Ix yawing moment due to nondimensional Cnp cross product of inertia, slug-ft 2 Ixz roll rate derivative, OCn/O(pb/2 V ) , rad -I pitch moment of inertia, slug-ft 2 Iy yaw moment of inertia, slug-ft 2 yawing moment due to nondimensional I z Cllr yaw rate derivative, _ Cn/O (rb/2 V), KEAS equivalent airspeed, knots rad 1 LASRE linear aerospike SR-71 experiment yawing moment due to angle of sideslip Cn 0 m mass, slugs derivative, OCn/O _ , deg "1 roll rate, rad/sec yawing moment due to aileron derivative, P C nsa _Cn/OSa , deg -1 roll acceleration, rad/sec 2 yawing moment due to rudder derivative, Cns, pitch rate, rad/sec q _Cn/_r , deg "1 dynamic pressure, lb/ft 2 normal force coefficient CN pitch acceleration, rad/sec 2 normal force bias, linear coefficient CN b estimate for tx = 0 ° r yaw rate, rad/sec yaw acceleration, rad/sec 2 normal force due to nondimensional pitch CNq rate derivative, OCN/O(qc/2 V), rad -1 ref aerodynamic moment derivatives corrected to the 25 percent c moment reference normal force due to angle of attack CN_ (F.S. 900) derivative, OCN/OO_, deg -1 S SR-71 reference area, 1605 ft 2 normal force due to elevon derivative, CN6, OCN/OSe, deg "1 T thrust, lb side force coefficient Cy U body x-axis wind-relative velocity, ft/sec side force bias, coefficient estimate for CY b body x-axis wind-relative acceleration, 13=0 o ft/sec 2 side force due to nondimensional roll rate Crp body y-axis wind-relative velocity, ft/sec derivative, 3Cr/O(pb/2V), rad "l body y-axis wind-relative acceleration, side force due to nondimensional yaw rate ft/sec 2 Cy, derivative, OCr/O(rb/2V), rad -1 V true airspeed, ft/sec side force due to angle of sideslip Cy_ flightpath wind-relative acceleration, derivative, _Cy/_, deg -I ft/sec 2 side force due to aileron derivative, Crsa W body z-axis wind-relative velocity, ft/sec _Cy/_a, deg "1 body z-axis wind-relative acceleration, side force due to rudder CY_r ft/sec 2 derivative,OCy/OSr, deg "1 wing-reference-plane angle of attack, deg D drag, lbf angle of sideslip, deg EGT exhaust gas temperature time rate of change of angle of sideslip, excess thrust, lbf F e rad/sec ES.
fuselage station, in.
_a aileron deflection, deg American Institute of Aeronautics and Astronautics _e elevon deflection, deg incorporated into a lifting body configuration, a cooperative project between the NASA Dryden Flight 8r rudder deflection, deg Research Center, Edwards, California, and Lockheed pitch angle, deg Martin was initiated. The project approach was to flight test an aerospike rocket using an SR-71 aircraft as the time rate of change of pitch attitude, carrier vehicle in a project known as the Linear rad/sec Aerospike SR-71 Experiment (LASRE). The primary ¢ roll angle, deg goal of the project was to gather installed rocket engine performance data at flight conditions approximating the time rate of change of roll angle, rad/sec X-33 trajectory.
Introduction The LASRE lifting body configuration is roughly a 20-percent scale model of an X-33. The entire test The goal to dramatically reduce the cost of access to apparatus, known as the LASRE pod, is shown mounted space has prompted NASA to fund the development of on the SR-71 aircraft in figure 2. The large size of the the X-33 Single Stage to Orbit Technology LASRE pod significantly altered the aerodynamics of Demonstrator Vehicle. Lockheed Martin Skunk Works, the SR-71 configuration and, therefore, the early flights Palmdale, California, has been chosen to build the of the program were used to clear the flight envelope X-33, which incorporates a linear aerospike rocket before attempts were made to carry volatile propellants engine (built by Boeing Rocketdyne, Canoga Park, and fire the rocket engine. These envelope expansion California) into a lifting body configuration. The flights were used to address flutter clearance, stability principle advantage of the aerospike rocket is the and control concerns, and to obtain performance data on inherent altitude compensation provided by the nozzle.
the LASRE configuration. Many of the concerns that This altitude-compensating ability theoretically allows required flight testing were identified during a series of the rocket engine to achieve increased specific impulse wind tunnel tests conducted prior to fabrication of the performance during the low-altitude portion of a flight LASRE pod 3. These concerns included such issues as as compared with a conventional bell nozzle rocket transonic pitch-trim authority and the large drag caused (fig. 1).
by the pod.
The aerospike rocket was first developed and ground This report presents the aerodynamic data that were tested in the 1960's i' 2, but has never been flight tested.
gathered during the envelope expansion flights These In order to obtain flight data on an aerospike rocket data include flight measured stability and control data I Aerospike nozzle -_ / ii/ i////// Specific _#:i. !__ _i_, impulse / / J'/ _--B , ell 4, .. ," nozzle s Bell nozzle Aerospike nozzle Sea level High altitude Altitude 960227 Figure 1. Qualitative comparison of aerospike and bell nozzle specific impulse.
American Institute of Aeronautics and Astronautics EC97 44295-103 Figure 2. LASRE configuration in flight.
and performance data up to Mach 1.8. The flight data on wind tunnel predictions of significant additional drag are compared with preflight wind tunnel predictions.
resulting from the LASRE pod. The thrust enhancement The importance of flight simulation to envelope included trimming the maximum rotor speed and core expansion testing is discussed in detail. Use of trade fuel flow to the top end of their operating bands and names or names of manufacturers in this document does uptrimming the turbine exhaust gas temperatures not constitute an official endorsement of such products (EGT). The pilot-controlled EGT uptrim was only used or manufacturers, either expressed or implied, by the during the transonic and low-supersonic acceleration.
National Aeronautics and Space Administration.
The aircraft internal structure was extensively modified to support mounting a structure of up to 14,500 Ib on the Configuration Description top of the fuselage. 4 Baseline SR-71 Aircraft LASRE Pod A Lockheed Martin SR-71A aircraft was used as the The LASRE components mounted to the top of the carrier vehicle for the LASRE. The SR-71A aircraft is a SR-71 were referred to as the canoe, kayak, reflection two-place, twin-engine aircraft capable of cruising at plane, and model (fig. 3). Collectively, these structural speeds up to Mach 3.2 and altitudes up to 85,000 ft.
components were referred to as the LASRE pod. The Twin all-moving rudders mounted on top of the engine canoe was installed on the SR-71 fuselage and was nacelles provided directional control, while inboard and designed to contain the gaseous hydrogen fuel and outboard elevons provided longitudinal and lateral liquid water needed for cooling. The kayak, located control. The inboard and outboard surfaces moved beneath the reflection plane and on top of the canoe, sets simultaneously. However, the outboard elevons were the model incidence angle to 2 ° nosedown to align the lower part of the model with the local flow over the top rigged with an additional 3 ° trailing-edge-up incidence of the SR-71 airplane. The reflection plane was mounted in comparison with the inboard elevons. The control on top of the kayak to help promote uniform flow in the surface actuators were powered using two independent region of the model. The model was designed to hydraulic fluid systems. Two Pratt & Whitney J58 approximate a half-span lifting body with a 20 ° turbojet engines were used to power the aircraft.
swept-cylinder leading edge and spherical nose. Liquid SR-71 Modifications oxygen and ignitor materials required to operate the rocket engine were stored in the model. The model was The SR-71 aircraft for the LASRE program included mounted vertically so that sideslip of the SR-71 airplane thrust enhancement and structural modifications. It was would impart angle of attack on the model. With a full decided to increase the thrust of the J58 engines based load of expendables the pod weighed approximately American Institute of Aeronautics and Astronautics
Canoe--_ !
I " I Kayak J F.S. 1355 -/ -- B.L. 340.2
c
980394 Figure 3. Side and planform views of the LASRE configuration.
14,140 lb. The total empty weight of the LASRE Accelerations were corrected to the c.g. using angular configuration was approximately 74,870 lb. Fuel loads rate information from the strapdown sensors. All control of up to 62,000 Ib have been used during the flight tests.
surface positions were measured with the exception of To compensate for center of gravity (c.g.) shifts caused the right outboard elevon. The inboard and left outboard by the pod weight, 5000 lb of fuel in the forward tank elevon actuators were instrumented with hydraulic pressure sensors for the trailing-edge-up deflection.
was considered unusable for the flight.
Vehicle weight and c.g. were obtained using fuel tank Instrumentation measurements.
The SR-71 was equipped with a complete set of air Analysis Methods data and inertial instrumentation. Free-stream pitot- static air data were obtained from a calibrated Stability and control and aircraft transonic noseboom. Angle of attack (_) and angle of sideslip (13) performance data were of the most interest during the data were obtained from a 4-hole hemispherical probe early LASRE test flights. The methods used to analyze dog-legged to the noseboom. The angle of attack is these data are now discussed, including the importance referenced to the wing reference plane which is of preflight wind tunnel predictions and flight 1.2 ° nosedown in incidence compared with the fuselage simulation.
centerline reference plane. Angle of attack and angle of sideslip measurements were lagged from 0.2 to Stability and Control 0.4 seconds because of the pneumatic plumbing. These lags were accounted for by time skews in the data Stability and control derivatives were obtained from analysis. Pitch and roll attitude information were flight data using maximum-likelihood parameter obtained from the SR-71 inertial navigation system. estimation techniques. 5'6 As is normally the case, Angular rate and linear accelerations were measured longitudinal and lateral-directional analysis were using strapdown sensors installed on the SR-71.
performed separately.
American Institute of Aeronautics and Astronautics Longitudinal lateral-directional stability and control derivatives. The state equations used in the lateral-directional analysis Pitch doublet maneuvers were flown at specified are: Mach numbers and altitudes to obtain the longitudinal stability and control derivatives. The state equations ?/S (6) used in the longitudinal analysis are: 13 = _--QCy + psinct - rcosa + gsin_cos0 ?/S pl x - i'lxz = ?tSbC l + qr(ly - lz) + pqlxz (7) 6c = -_--QCNCOSa + q - tanl](pcosa + rsintx) (1) + g(cos¢cos0cosa + sin0sinct) i'l z - Plxz = ?ISbC n + pq(l x - ly) - qrlxz (8) v t_ = p + qtan0sin¢ + rtan0cos¢ (9) (1 = ?ISCCm + rp(l z- I x) + (r 2 - p2)lxz (2) The reference span, b, was 56.7 ft. The response 0 = qcost_-rsin¢ (3) parameters measured and estimated were angle of sideslip, roll rate, yaw rate, bank angle, and lateral The response parameters measured in flight and acceleration. Angle of sideslip and lateral acceleration compared with estimations were angle of attack, pitch measurements were corrected to the c.g. Because the rate, pitch attitude, and normal acceleration. Angle of angle of sideslip calibration was suspect, lateral attack and normal acceleration measurements were acceleration was weighted heavier in the analysis than corrected to the c.g. Normal acceleration was weighted angle of sideslip. The force and moment coefficient heavier in the analysis than angle of attack because the equations are: angle of attack calibration was suspect. The force and moment coefficients were expanded using the linear b approximation: Cy = CYb + Cyfj_ + "_-_( Cy pp + CYrr ) (10) c + Cya Sa + Cys _r C N = CNb + CNa + _'-_CNq q + CN68e (4) c Cl = Clb + CI_ + 2_(ClpP + Ctrr) (11) C m = Cmb + CraaO[ -k _--_Cmq q + Crn8Se (5) + Cl_Sa + Ci6rSr The coefficients are based on a reference area of 1605 ft 2 and a mean aerodynamic chord, c, of 37.7 ft. b C n = Cnb + Cn_ _ + _-Q(C%p + Cnr ) The moment reference is at 25 percent c, which is at (12) fuselage station (ES.) 900. The coefficient with the + CnsSa + Cn_Sr subscript b is a linear extrapolation from the coefficient at the average angle of attack of the maneuver to zero The coefficient with the subscript b is the value of the angle of attack. 5 Axial force coefficients were not used coefficient at zero angle of sideslip. All the lateral- in this analysis because the axial force derivatives were directional derivatives in equations 10--12 were not expected to affect flying qualities and because it is estimated in the analysis. Only Cn , C 1 , C , and . . 13 t_ nsr generally difficult to get good identifiability of these CI_,, results are presented m this report because these derivatives. Axial loads, however, were of importance to are the most relevant to the configuration stability the performance analysis and are discussed later. All of and control.
the longitudinal derivatives in equations 4 and 5 were Performaq¢¢ estimated in the analysis. Only Cm, _ and Cms" results are presented in this report because these are of crucial The SR-71's J58 engines were not instrumented for interest to the configuration stability and control.
thrust measurements. Consequently, thrust and drag could not be independently identified and an excess -hm.tal:D_imc, am_ thrust performance analysis was used instead. Excess Yaw and roll doublet maneuvers were flown at thrust was obtained from flight data and simulator data specified Mach numbers and altitudes to obtain the and then the two were compared.
American Institute of Aeronautics and Astronautics Acceleration maneuvers at low angle of attack and Simulation near zero angle of sideslip were used to obtain the Flight simulation was used extensively in preparation performance data. Excess thrust, Fe, is defined as: 5 for the LASRE flight tests. Two simulators were used; a workstation-based batch simulator and a piloted real- F e = (TcosGtcos13- D) time full-cockpit simulator. Wind-tunnel determined stability and control increments and drag increments cos ¢ cos 0 sin acos 131 (13) resulting from the LASRE pod were added to the =m(Z-mg[ +sin¢cosOsin13 [ baseline SR-71 aerodynamic model 7 for use in the L - sinOcos(xcosl3.J simulations. The batch simulation was used to obtain stability and control derivatives, trim elevon, and hinge where moments predictions for the LASRE configuration. The real-time simulation was used for pilot training and performance estimation.
_, _ uft + vf_ + ww V Results and Discussion t_ = ax-qW+rv-gsinO (14) This section of the paper focuses on stability f_ = ay-ru+ pw+gsin_cosO and control and performance results. Results are = a z-pv + qu +gcos¢cosO discussed from baseline SR-71 flight tests without the LASRE pod installed, wind tunnel tests, flight Wind Tunnel simulation studies, and flight test with the LASRE pod installed.
Before fabrication of the LASRE pod, wind tunnel testing of a 4-percent scale model was completed to Baseline SR-71 Flight Results obtain stability and control increments and drag Stability and Control increments resulting from the pod. 3 Wind tunnel data were obtained for the baseline SR-71 and the In an effort to verify the existing SR-71 aerodynamic LASRE configuration. Increments were determined by model, 7 a series of stability and control doublet subtracting the baseline SR-71 wind tunnel data from the maneuvers were flown and analyzed for the baseline LASRE configuration wind tunnel data.
SR-71 aircraft. Figure 4 shows the Mach and altitude 70 x 103 KEAS 0 Pitch doublets / I-I Yaw-roll doublets i 6O -- LASRE flight envelope i .... _._eo_.._..
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I0 _ I t t i ! i / ¢ 0 .5 1.0 1.5 2.0 Mach 980395 Figure 4. Flight conditions for baseline SR-71 stability and control test points.
American Institute of Aeronautics and Astronautics envelope for the LASRE configuration and the test simulation data. Figure 6 shows the elevon effectiveness, conditions for the baseline SR-71 aircraft stability and Cm6 ' . Good agreement is observed between the flight control maneuvers shown in this report. The and simulation data except at subsonic Mach numbers flight-derived stability and control derivatives for the where the flight-derived effectiveness is as much as baseline aircraft were compared with estimates from the batch simulator. 20-percent less than the simulation results.
Longitudinal Stability and Cotltrol Lateral-Directional Stabili_ and Control The longitudinal stability derivative, Cm, and The lateral-directional stability and control derivatives the elevon effectiveness derivative, Crns, for the for the baseline SR-71 aircraft are shown in figures 7-10.
baseline SR-71 aircraft are shown in figures 5 and 6, Cn_ and Cns r have been corrected to the moment respectively. Data were obtained at c.g.'s ranging from reference. The directional stability derivative, Cn, 19-24 percent c. The data in figures 5 and 6 were shown in figure 7 shows slightly less stability than corrected to the moment reference used in the simulation simulation from moderate subsonic to sonic Mach which is at 25 percent c (SR-71 F.S. 900). The circles in numbers. The dihedral effect, C/l_, (fig. 8) shows figure 5 represent flight-derived longitudinal stability reasonably good agreement with the largest deviation derivatives and the squares represent simulation results around Mach 1.2, where the flight data showed less for the same flight conditions. The aircraft is fairly stability. Figure 9 shows the rudder control effectiveness, flexible, 7 and therefore some of the variability in the data Cn_ r, which agrees well with simulation, except is a result of test points at slightly different dynamic subsonically where the flight-derived effectiveness is pressures. The solid line in figure 5 represents a hand approximately 15-percent less than simulation. The fairing of the flight data using Cramrr-Rao bounds 5 as an indication of the maneuver quality. The dashed line aileron control effectiveness, Ct_, shown in figure 10, represents a fit of the simulation data. As can be seen, agreed well with simulation except for slight differences there is fairly good agreement between the flight and at high subsonic Mach numbers.
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0 .5 1.0 1.5 2.0 2.5 Mach 980396 Figure 5. Baseline SR-71 longitudinal stability derivative corrected to the moment reference.
American Institute of Aeronautics and Astronautics _O_ Flight i i¸ --D- Simulation - .001 - .002 - .003 Cm_)ersf' - .004 deg -1 - .005 - .006 LI i i - .007 - .008 0 .5 1.0 1.5 2.0 2.5 Mach 980397 Figure 6. Baseline SR-71 elevon effectiveness derivative corrected to the moment reference.
.0025 Flight 0 i ---E}- Simulation .0020 ................................................. /8\ .................................................................................
.0015 ................ _'"i ................... _ !Direction of Cnl3ref' i v-v- i _\ increased i i _ i stability .oolo ................ T !..................................... " ....... ' ...................................
.0005 i 0 .5 1.0 1.5 2.0 2.5 Mach 980398 Figure 7. Baseline SR-71 directional stability derivative corrected to the moment reference.
American Institute of Aeronautics and Astronautics
I i --O--iF,ight
I i --D-i Simulation -.ooo, I........................................ i................................................................................... i........................................
- .0010 _ ....
-.ools ........................................................................... i......................................... i.......................................... .........................................
clly deg -1 I T Direction of -.0020 ............................. 14 .......................................................................................................................... i. increased / i stability I -.0025 -.0030 ............................ I .................................................... T ........................................ _ ......................................... i.......................................
- .0035 0 .5 1.0 1.5 2.0 2.5 Mach 980399 Figure 8. Baseline SR-71 dihedral effect.
i ---O--! Flight -.-0- i Simulation
o r i
- .0005 .................................................................................................................................................................... i.......................................
CmSrref' - .0010 deg -1 -.0015 - .0020 0 .5 1.0 1.5 2.0 2.5 Mach 98o4oo Figure 9. Baseline SR-71 rudder effectiveness derivative corrected to the moment reference.
1o American Institute of Aeronautics and Astronautics .0025 .0020
io
.0015 ClSa, deg-1 .OOLO .o005 0 .5 1.0 1.5 2.0 2.5 Mach _0401 Figure 10. Baseline SR-71 aileron effectiveness derivative.
Performance transonically. 3 For 4 ° _, typical trim _ for the SR-71, the wind tunnel test predicted a maximum noseup The performance of the J58 engine is strongly linked increment at Mach 1.05 and a significant nosedown to ambient air temperature. Baseline SR-71 flight data increment at Mach 1.2 (fig. 12).
collected from different days with different ambient temperatures were analyzed in order to evaluate the Performance excess thrust model in the simulator. Excess thrust Figure 13 shows the wind-tunnel predicted trimmed performance from the flight data was obtained using equation 13 and was compared with the simulations for drag for the LASRE configuration plotted with the the same flight conditions. In all cases, the simulator baseline SR-71 trimmed drag for 4 ° _. As observed, wind tunnel results showed that the addition of the overpredicted the excess thrust for Mach numbers between 0.95 and 1.2; in some cases by as much as LASRE pod resulted in a maximum drag rise of nearly 5000 lb. Two full-afterburner level accelerations at an 70 percent.
altitude of 30,000 fi are shown in figure 11. In one case Simulation Studies the ambient temperature was 6 °C above the standard day temperature and in the other case it was 3.6 °C Stability and Control below the standard day temperature. As observed, the simulator overpredicted excess thrust in both cases, and Simulations showed that obtaining trimmed flight more so for the wanner day. Unfortunately, it was required elevator deflections of approximately 5 ° not practical to improve the simulator performance noseup at Mach 1.2 at the altitudes used for transonic fidelity because of the complexity of the aerodynamic acceleration. The simulator showed that this trim and propulsion models combined with the scarcity of requirement approached the maximum capability of the flight-to-simulation performance comparisons.
elevon actuator power if one of the two SR-71 hydraulic systems were to fail.
LASRE Pod Wind Tunnel Predictions Performance Stability and Control The LASRE configuration performance capability Wind tunnel tests predicted that the LASRE pod would cause a significant change in pitching moment was analyzed by wings-level accelerations performed in AmericanInstitute of Aeronautics and Astronautics 18 x 103
+ +
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F e , ; Ib 8 -- ! ....... +. ...... 4- --_ . --+ .95 1.00 1.05 1.10 1.15 1.20 Mach 980402 (a) Ambient temperature 6 °C greater than standard day temperature.
18 x 10 3 i 14 _ ,, • JI I s **_" . ._+.++_ !¢" !,, , _- Simulation : ,---" 12 ........................ l ................. _ _!_ +Z\ ..................................................... I ................................. / ..... i...............................................
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I .95 1.00 1.05 1.10 1.15 1.20 Mach 980403 (b) Ambient temperature 3.6 °C less than standard day temperature.
Figure 11. Flight and simulation excess thrust results for the baseline SR-71 aircraft during level accelerations at 30,000 ft altitude.
American Institute of Aeronautics and Astronautics
.005
C m increment - .005 - .010 0 .5 1.0 1.5 2.0 Mach 980404 Figure 12. Predicted C m increment resulting from the LASRE pod at 4 ° o_.
.040 .035 .030 Drag coefficient .025 .020 .015 I .010 .8 .9 1.0 1.1 1.2 1.3 1.4 1.5 Mach 980405 Figure 13. Predicted trimmed drag of the LASRE configuration compared to the baseline SR-71.
American Institute of Aeronautics and Astronautics the simulator using the standard day atmospheric the aerospike rocket in flight. Flutter clearance consisted temperature profile. 8 Under standard day conditions, the of a series of longitudinal pitch pulses during level LASRE configuration was predicted to be capable of accelerations or constant KEAS climbs. This report will obtaining the original project requirements of Mach 3.2 not discuss the flutter clearance except to say that the test points. The performance pinch point is during required flight envelope was successfully cleared for transonic acceleration where excess thrust is at a flutter with no concerns. Stability and control envelope minimum. The addition of the LASRE pod added as expansion included a series of pitch doublets and yaw- roll doublets. 5 In some cases the pilot would stabilize much as 70-percent more drag transonically than for the the aircraft at a specified Mach and altitude and perform baseline SR-71. Figure 14 shows a simulation of fuel a series of doublets. In other cases, the pilot would usage during a 25,000 ft. altitude level acceleration to perform a single doublet at a specified Mach number 450 knots equivalent airspeed (KEAS) (Mach 1.12) during an acceleration or deceleration. In all cases, these followed by a constant KEAS climb to 31,800 ft (Mach doublets demonstrated the acceptability of the LASRE 1.3). Results were obtained for a range of temperatures configuration handling qualities in real-time. The between 10 °C warmer than a standard day and 10 °C maneuvers were analyzed postflight to obtain stability colder than a standard day. As can be seen, an additional and control derivatives for the LASRE configuration.
11,000 lb. of fuel was required on the +10 °C-day as Because of structural concerns associated with the pod, compared to a standard day. As discussed previously, the doublet size was limited by angular acceleration to the transonic performance simulation was already be less than 8°/see 2 pitch acceleration, 4.5°/see 2 yaw suspect (fig. 11). Given the uncertainties in the acceleration, and 43°/see 2 roll acceleration.
performance simulation for the baseline SR-71 and the drag predictions for the addition of the LASRE pod, it The simulation-predicted transonic noseup pitch trim was again left for flight test to provide the definitive requirement was shown to approach the limit of elevon performance answers.
actuator power available if one of the two SR-71 hydraulic systems should fail. This potentially Flight Envelope Expansion dangerous situation required limiting the aircraft speed Both flutter envelope expansion and stability and and e.g. envelopes until flight envelope expansion control envelope expansion were required prior to firing determined the actual pitching moments. The flight 55 x 103 5o ...................... i............................ ............................
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........................... i....................................................... i.......................... _ ................... _i -''-'x' ............. i................... _-_-_. ...................
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! ; i i i _ i i 11,000 Ib fuel 35.......................... i........................... ; ........................... i.......................... i......................................... \ +io cl .....
.90 .95 1.00 1.05 1.10 1.15 1.20 1.25 1.30 Mach 9eo4oe Figure 14. Simulation results of the performance effects as a result of ambient temperature during a transonic acceleration and climb to Mach 1.3.
American Institute of Aeronautics and Astronautics envelope was limited to avoid a flight configuration that Again, these results were corrected to the 25 percent c could not recover from a failure of one of the two moment reference location. The circles represent flight hydraulic systems. For this reason, instrumentation was determined derivatives and the squares represent added to the inboard elevons and to the left outboard derivatives obtained from the batch simulation. The solid elevon to measure hydraulic actuator pressures for the line is the Cramdr-Rao-based hand fairing of the flight- elevon trailing-edge-up deflections. Real-time flight monitoring of these hydraulic pressures was done to determined derivatives for the test points shown in ensure that the aircraft would not fly into a flight regime figure 15 that are not rocket test points. The dashed line where it could become hinge-moment limited in the is the fit of the simulation predictions for the LASRE event of one hydraulic system failure.
derivatives at the same test points. The solid symbols LASRE Configuration Stability and Control Flight represent the flight and simulation derivatives for the Results lower KEAS rocket test points shown in figure 15. As The stability and control test points presented in this can be seen in figure 16, flight determined values of Cm, _ report for the LASRE configuration are shown in agreed fairly well with the simulation predictions.
figure 15. It should be noted that the pitch and roll Supersonically, flight data at the lower KEAS rocket test stability and control derivatives are a function of the points showed the same derivative values as the higher equivalent airspeed because of the flexibility of the KEAS data, whereas the simulations indicated that lower SR-71 aircraft. 7 The test points represented by the open symbols in figure 15 were chosen so that the derivative KEAS effects would improve the stability. As observed results could be plotted as a function of Mach number in figure 17, at transonic Mach numbers, flight derived without confusion caused by the flexibility effects. The Cm8" agreed fairly well with the predictions.
solid points correspond to lower KEAS flight conditions Subsonically, the elevon effectiveness is as much as at which the aerospike rocket test firings are planned.
20-percent less than the predictions. Comparing Longitudinal Stability and Control figure 17 with figure 6, the aerodynamic model ofelevon effectiveness for the baseline SR-71 was also in error The longitudinal stability and elevon control effectiveness derivatives are shown in figures 16 and 17. subsonically at approximately the same magnitude.
KEAS 50 103 0350 400 450 0 Pitch doublets / r" ,•" j r-i Yaw-roll doublets / ," , • • Doublets at rocket test points / •• i•" _" • • j LASRE fl,ght envelope _ •• , r_r "r 40 .................................... ,................. • ...............................................
....... .........................
i i • • i js S
' ./;;7
3O i • r 7 ........................ i .................................................
20 ,0 , ................................................
/ i _' IO, J ,: i s • s i i ,,/,: i t • I i i I _ s ,'// _ lO i ;' _ .................
/1"1 :: ' '/' i t t t i I' i , , s i i s I s !a I t !
il I 0 .5 1.0 1.5 2.0 Mach 980407 Figure 15. Flight conditions for LASRE configuration stability and control test points.
American Institute of Aeronautics and Astronautics .0O4 --O-- Flight i _' ---E}- Simulation i • Flight (low KEAS) i • Simulation (low KEAS) .002 ................. ___._!6, ................. T .......................... i......................... _T ......................... !! ...................... D_,ectionl ......................... of ! i u_-_- ii 1 increased ! i _ i stability deg-1 .................................................. i..........................
- .002 .................................. ,,_--,-,-_ ......
i '
- .004 ................................................................................................................................................................
i i - .006 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 9804O8 Figure 16. LASRE longitudinal stability derivative corrected to the moment reference.
O O i • J : --O--- Flight --E}- Simulation ..........
• ! Flight (low KEAS) • i Simulation (low KEAS) .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980409 Figure 17. LASRE elevon effectiveness derivative corrected to the moment reference.
American Institute of Aeronautics and Astronautics Pitch Trim given flight condition. The results depended on the elevon control effectiveness, Cm_ e , and the longitudinal Transonic pitch trim authority was a concern because stability, Cm. Based on figure 16, Cm_ from flight of wind tunnel predictions. A comparison between agreed well with predictions. Figure 17 showed that simulation prediction and flight measured trim elevon is shown in figure 18 using data from two different flights. Cms _ agreed well, except at subsonic speed. Based on The actual amount of trim required is a function of e.g.
figure 17, the LASRE simulator value for Cms _ was and therefore varies slightly between flights when plotted incremented as a function of Mach according to the against Mach number. The data points plotted are for following table: wings-level flight with a normal load of approximately 1-g. The flight data were low-pass filtered at 0.1 rad/sec Table 1. Flight determined Cm_" increment added to to remove the transient effects. In general, more the simulation.
nosedown trim was required than had been predicted throughout the Mach range. The trim agreed fairly well Mach 0 0.68 0.9 3.2 in the Mach 1.05 to 1.15 range. The largest change was at Mach 0.9. At this speed close to 1 ° nosedown trim was Cms _ increment 0.0014 0.0014 0.0000 0.0000 required compared with a 2 ° noseup trim requirement that had been predicted. Also, at Mach 1.2, 2 ° less noseup trim was required. Since the transonic estimates Using the batch simulator, the pitching moment of C m and Cmo (figs. 16 and 17) agreed fairly well increment caused by the LASRE pod was obtained from with pr_edictions,°t]ae differences seen in transonic elevon flight and was compared with the preflight prediction in deflections with simulator predictions were solely a figure 19. There are two regions of significant change.
result of mispredicting the zero-lift pitching moment The region from M = 0.9 to M = 1 shows a considerably increment in the wind tunnel test.
larger pitch-up increment and the region around M = 1.2 shows no pitch-down increment.
Pitching Moment A ctuator Pressu res The batch simulator was used to determine the increment in Cm that was required to make the As a result of concerns in reaching hinge-moment simulation elevon deflection match the flight value at a limits during transonic acceleration, elevon hydraulic Aircraft nosedown O Flight Simulation o o Trim elevon, deg O O -2 O o Oo -4 Aircraft noseup -6 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980410 Figure 18. Elevon trim comparison between flight and simulation predicted results.
American Institute of Aeronautics and Astronautics .oio .oo5 i i i i Cm increment I / li // - .OOS .................. _ .................... ;.................... i............................. _...-* .................... _................... .............................................................
!
F.ght i Z , ---.. I I - .010 i i Predi.ted 0 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 2.0 Mach 980411 Figure 19. Pitching moment increment as a result of the LASRE pod, obtained from flight data and wind tunnel predictions.
actuator pressures were measured and observed during stability, C n , and dihedral effect, CI , were less stable flight. Figure 20 shows the outboard and inboard than the simulator predictions had indicated. Figure 21 hydraulic pressure measurements during wings-level, shows Cn as much as 30-percent less than predicted.
l-g trim conditions. With zero hinge moment, the As discussed in reference 7, there were no flexibility pressure reads 1000 lb/in 2. The upper limit for safe corrections for the yaw axis. The flight data verified this, operation was 2100 lb/in 2. As can be seen in as there was little difference between the data at the low figures 20(a) and 20(b), the limit was not reached by KEAS test points and at elevated KEAS test points either the outboard or inboard elevons. The data below (with the exception of one maneuver at Mach 0.9).
Mach 1 and above Mach 1.2 show that the actuator Figure 22 shows C t as much as 50-percent less stable pressures were lower than predicted, which is consistent than predictions. C-'_omparisons with figures 7 and 8 with the reduced noseup trim requirement (fig. 18). The outboard elevon pressure did indicate that with one showed that similar trends in misprediction were seen in hydraulic failure the elevon would be hinge-moment the baseline SR-71 aerodynamic model.
limited at Mach 1.0 to 1.04 in nosedown capability (i.e.
The rudder and aileron effectiveness control the pressure was less than the 500 lb/in 2 minimum criteria). This was not a safety concern for two reasons: derivatives are shown in figures 23 and 24. The rudder (1) the inboard elevons were not hinge-moment limited effectiveness, Cnrr, was less than predicted below and therefore trim authority still existed, and (2) if both Mach 1.3 and the aileron effectiveness, Cl_a, showed inboard and outboard elevons did become hinge- good agreement. Subsonically, the rudder effectiveness moment limited the aircraft would pitch up slowly and misprediction was of the same magnitude as the decelerate to a point at which the control authority misprediction of the baseline SR-71 rudder effectiveness would return.
(fig. 9). However at the low supersonic Mach numbers Lateral-Directional Stability and Control the misprediction in Cns r is caused by the misprediction of the pod effect from the wind tunnel. At worst, the The angle of sideslip derivatives are shown in rudder effectiveness was 23-percent less than predicted figures 21 and 22. For the most part both the directional at Mach 1.05.
American Institute of Aeronautics and Astronautics 0 Flight Simulation Trailing-edge-up limit 2O00 Actuator hydraulic pressure, Ib/ln 2 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980412 (a) Outboard elevon.
FligLt n ;i o Simulation Trailing-edge-up limit Actuator hydraulic pressure, Ib/Jn2 I 51111 i ! Trailing-edge-down limit !
.6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980413 (b) Inboard elevon.
Figure 20. Elevon hydraulic pressure.
American Institute of Aeronautics and Astronautics .0025
I i -o- Right
i. i --El'- Simulation _ • Flight (low KEAS) i , • Simulation (low KEAS) .0O2O o_o ,, _._ 8- o, .OO15
................ °N__-i__i'_ ' '-'x-' ...................... _. :t
Cn _ref' deg -1
i °1
........................ i ....... _ . t i _' , .OOlO .OOO5 Direction of _- increased stability .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980414 Figure 21. LASRE directional stability derivative corrected to the moment reference.
i - .0010
i
Direction of - .OO15 J'""','i ......................... i ..................................................................................... increased stability
/
i I I - .OO20 ...........................................
' ]
! --O-- Flight i i i_m Simulation I - _ • Flight (low KEAS)' • Simulation (low KEAS) 1.0 1.2 1.4 1.6 1.8 Math 980415 Figure 22. LASRE dihedral effect.
2O American Institute of Aeronautics and Astronautics .4 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980416 Figure 23. LASRE rudder effectiveness derivative corrected to the moment reference.
.0025 I_ _ E I _ _ F ight I _ i --El.- Simulation I _ _ i • FIght (low KEAS) .0020 1.................... _'_i ...................... i........... HI. Li.. Sl.mulat.iO.n(,owT KEAS) Cl_a, deg -1 .2 .4 .6 .8 1.0 1.2 1.4 1.6 1.8 Mach 980417 Figure 24. LASRE aileron effectiveness derivative.
American Institute of Aeronautics and Astronautics Simulation Update The flight data showed that the LASRE configuration accelerated at a rate significantly worse than simulation Flight determined corrections to C m , Cm8 ' , C%, results had predicted (fig. 25). The simulation, which C l_ , and Cns" were input into the piloted simulation for took into account the flight day temperature and the handling qualities evaluations. These evaluations flight-derived C m increments presented in figure 19, included aerospike rocket firings at the specified test underestimated the required fuel usage by 10,000 lb. To compensate for this discrepancy, an effort was made to conditions and emergency situations such as engine and increase the fidelity of the simulator. The excess thrust hydraulic system failures. In all cases, the simulations computed from flight and simulator data is shown in showed acceptable handling qualities and that the figure 26. The excess thrust increment between flight and aircraft responses were within acceptable load factor simulator data was compared with the predicted trimmed and angle of sideslip limits.
drag increment caused by the pod (fig. 13) in figure 27.
The excess thrust increment was equivalent to a LASRE Configuration Performance Flight ReSORs 40-percent pod drag increase at Mach 0.98 and approximately 25-percent pod drag increase at The most critical part of the flight envelope pertaining Mach 1.02 and higher. This excess thrust increment was to performance was the transonic acceleration. The subtracted from the simulator performance model and LASRE configuration had significant excess thrust at all the acceleration was simulated using the same ground Mach numbers except at transonic conditions. The first track as the flight data. As shown in figure 28, the LASRE flight occurred at the end of October on an simulator results now agreed well with the flight results.
unseasonably warm day. A 27,000 ft altitude level acceleration was used to accelerate the aircraft through The second LASRE flight occurred on a nearly the transonic Mach numbers while simultaneously standard temperature day. A level acceleration was obtaining flutter as well as stability and control tesl data.
performed at an altitude of 25,000 ft where the ambient The free-stream temperature at 27,000 ft was 9 °C temperature was just 1 °C above the standard day warmer than a standard day. As a result of the hot temperature. The fuel performance plot is shown in temperature, excess thrust was so low that the aircraft figure 29 and compared with simulation results for was only able to accelerate to Mach 1.17 before reaching the fuel-low limit. accelerations at +4 °C and +7 °C. The simulation results 46 x 103 Fuel, 34 Ib .90 .95 1.00 1.05 1.10 1.15 1.20 Mach 980418 Figure 25. Fuel performance for a transonic acceleration at 27,000 ft.
American Institute of Aeronautics and Astronautics 5OO0 / / Excess / / thrust, Ib IOO0 - 1000 .90 .95 1.00 1.05 1.10 1.15 1.20 Mach 980419 Figure 26. Flight and simulator excess thrust results for transonic acceleration at an altitude of 27,000 ft.
.014 --0_ Predicted drag increment resulting from the pod (4 ° 0{) .012 ---D-- Excess thrust decrement for 27,000 ft transonic ,010 acceleration .008 Drag .O06 and Fe coefficient .O04 .002 - .002 .6 .7 .8 .9 1.0 1.1 1.2 1.3 1.4 Mach 980420 Figure 27. Comparison of the excess thrust increment for the transonic acceleration at an altitude of 27,000 ft with the wind-tunnel predicted LASRE pod trimmed drag increment.
American Institute of Aeronautics and Astronautics 46 x 103 Fuel, 34 Ib .05 1.00 1.05 1.10 1.15 1.20 Mach 980421 Figure 28. Flight and simulator fuel performance results for the transonic acceleration at an altitude of 27,000 ft with the simulator excess thrust decremented according to flight results.
58 x 103 ..._ .. i i i ! _ 'Flight data i J <_N_ i [ Simulation data + 7 °C day • _,_ -& - _ i Simulation data +4°C day ] !
_k \! .................... l I ,. i , i Fuel, 44 Ib 3O 1.00 1.05 1.10 1.15 1.20 1.25 1.30 1.35 1.40 1.45 1.50 Mach 980422 Figure 29. Fuel performance for a transonic acceleration at an altitude of 25,000 ft and 455 KEAS climb to Mach 1.5.
American Institute of Aeronautics and Astronautics included the excess thrust decrement obtained from the 1 percent farther forward than on flight 3. Postflight simulations showed that the forward c.g. would account first flight (fig. 27). The simulator once again for less than 400 lb more fuel usage during the overpredicted transonic performance. Because drag is acceleration. This comparison further demonstrates that not a strong function of ambient temperature, it was small changes in temperature have a significant effect on concluded that the simulator does not correctly model transonic performance for the LASRE configuration.
J58 thrust changes as a function of ambient temperature.
The shape of the flight curve agreed fairly well with the Concluding Remarks +4 °C curve. After more simulation studies, it was concluded that the flight data for a +1 °C day could be Flight stability and control tests and performance tests approximated very well with a +5 °C day simulation.
have been successfully completed for the LASRE configuration at speeds up to Mach 1.8. Flight data have After completion of the initial level-altitude transonic been compared with preflight wind tunnel predictions.
accelerations that were required for flutter clearance, a more efficient piloting technique was used to improve Flight parameter estimation analyses of the LASRE the transonic penetration. Instead of accelerating at a configuration showed good agreement with wind tunnel level altitude, the acceleration began at an altitude of 28,000 ft and the pilot put the aircraft into a slight dive predictions of longitudinal stability, Cm, _, and elevon to help get through the transonic drag rise. The pilot control effectiveness, Cm8 ,, for transonic and leveled the aircraft at approximately Mach 1.07 and supersonic Mach numbers. Below Mach 0.9, the elevon 25,000 ft, which was the minimum altitude limit for control effectiveness was less than predicted by up to transonic Mach numbers. Results from two of these 20 percent. However, flight data shows that the baseline maneuvers flown on two different days are shown in SR-71 aerodynamic model also overpredicted elevon figure 30. During the acceleration from Mach 1.0 to 1.3, effectiveness by about the same amount. LASRE the flight 5 maneuver required approximately 5000 lb configuration transonic directional stability, C,_, was more fuel than the flight 3 maneuver. Two factors less than predicted with the worst case being 30-percent contributed to this poorer performance; slightly warmer less. As a result of the pod, the aircraft dihedral effect, temperature and the c.g. on flight 5 was approximately 55 x 103 Fuel, Ib 4O i .................................................................... i......................................................................... T ................................... ......
! i • i i i - i 3O 1.00 1.05 1.10 1.15 1.20 1.25 1.30 Mach 980424 (a) Available fuel.
Figure 30. Transonic accelerations from two flight days with different ambient temperatures.
American Institute of Aeronautics and Astronautics
i F i !l
• : _.° E _ _ i Jot| | Temperature deviation from standard day, °C -2 -4 1.00 1.05 1.10 1.15 1.20 1.25 1.30 Mach 980425 (b) Temperature deviation from standard day.
Cogog percent C ..-"__ ! .°.t* I Flight 5 i i
=o i i i
1.00 1.05 1.10 1.15 1.20 1.25 1,30 Math 980423 (c) Center of gravity.
Figure 30. Concluded.
American Institute of Aeronautics and Astronautics to answer is how much of the measured excess thrust C l , was also less than predicted by up to 50 percent.
difference is a result of wind tunnel misprediction of the Rudder effectiveness, Cn6 r , was as much as 23-percent LASRE drag increment and how much is caused by the less than predicted, whereas aileron effectiveness, CI_ ° , J58 thrust modeling error of the simulation. The flight agreed well with predictions. In general, simulator results can only conclusively state that the wind tunnel values for subsonic values of the baseline SR-71 prediction of the LASRE drag increment was within aerodynamic derivatives were not as good as the 25 percent for Mach numbers greater than 1.02.
supersonic values. Although LASRE stability and control effectiveness were lower than predicted in many References cases, the aircraft flying qualities were never lMartinez, A., Aerodynamic Nozzle Stud3; vol. 1, significantly worse than predicted.
Rocketdyne Final Report R-6582, North American Wind tunnel data had predicted that the pod would Aviation, Inc., Los Angeles, California, July 15, 1966.
cause a significant amount of noseup and nosedown 2Martinez, A., Aerodynamic Nozzle Study, Slipstream pitching moment as the aircraft accelerated through the Studies, vol. III, Rocketdyne Interim Report R-6273, transonic Mach numbers. The flight data showed North American Aviation, Inc., Los Angeles, California, significantly more noseup pitching moment increment July 31, 1965.
resulting from the pod than had been predicted, with the peak at Mach 0.95. The predicted nosedown pitching 3Moes, Timothy R., Brent R. Cobleigh, Timothy R.
moment increment at Mach 1.2 never materialized.
Conners, Timothy H. Cox, Stephen C. Smith, and Norm Shirakata, Wind-Tunnel Development of an SR-71 Piloted simulations using the flight-corrected stability Aerospike Rocket Flight Test Configuration, AIAA 96- and control derivatives were done for potential 2409, NASA TM-4749, June 1996.
emergency situations and aerospike rocket firings. In all cases, these simulations showed acceptable handling 4Corda, Stephen, David P. Lux, Edward T. Schneider, qualities and aircraft responses that were within and Robert R. Meyer, Jr., "Blackbird puts LASRE to the acceptable load factor and angle of sideslip limits.
Test," Aerospace America, vol. 36, no. 2, pp. 24-29, Feb. 1998.
The performance analysis of the LASRE configuration is still ongoing. Flight data clearly showed 5Maine, Richard E. and Kenneth W. Iliff, Application that the ability of the LASRE configuration to obtain of Parameter Estimation to Aircraft Stability and required project test points was highly dependent on Control - The Output-Error Approach, NASA RP-1168, ambient temperature because of the effect of June 1986.
temperature on J58 engine performance. The colder the temperature at altitude, the better the aircraft was able to 6Murray, James E. and Richard E. Maine, pEst accelerate to supersonic flight. It was also determined Version 2.1 User's Manual, NASA TM-88280, that the piloted simulator did not correctly model Sept. 1987.
ambient temperature effects on J58 engine thrust.
The simulator consistently overpredicted transonic 7Meyer, J. E., J. R. McMaster, and R. L. Moody, performance for the baseline SR-71 and the LASRE Handling Qualities of the SR-71, Lockheed Aircraft configuration. An excess thrust analysis was done to Corporation, Report no., SP-508, Burbank, California, quantify the performance difference between flight Oct. 29, 1964.
data and simulation results. On a LASRE flight with 8U.S. Standard Atmosphere, National Aeronautics ambient temperature approximately 9 °C above standard and Space Administration, United States Air Force, temperature, the excess thrust difference between flight United States Weather Bureau, ICAO Standard and simulation was 25 percent of the expected pod drag Atmosphere to 20 Kilometers, 1962.
increase at Mach 1.02 and greater. The difficult question American Institute of Aeronautics and Astronautics
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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORTTYPE AND DATES COVERED August 1998 Technical Memorandum 4.TITLE AND SUBTITLE 5. FUNDING NUMBERS Flight Stability and Control and Performance Results from the Linear Aerospike SR-71 Experiment (LASRE) WU 244-33-02-00-23-00-T15 e. _rrHOR(S) Timothy R. Moes, Brent R. Cobleigh, Timothy H. Cox, Timothy R.
Conners, Kenneth W. Iliff, and Bruce G. Powers 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER NASA Dryden Flight Research Center P.O. Box 273 H-2276 Edwards, California 93523-0273 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration NASA/TM-1998-206565 Washington, DC 20546-0001 11. SUPPLEMENTARY NOTES Presented at the AIAA Atmospheric Flight Mechanics Conference, August 10-12, 1998, Boston, MA. Timothy Moes, Brent Cobleigh, Timothy Cox, Timothy Conners, and Kenneth lliff, NASA Dryden Flight Research Center, Edwards, CA; and Bruce Powers, Analytical Services and Materials, Inc., Edwards, CA.
12a. DiST_iBUTION/AVAILABILITY STATEMENT 12b. DISTRIBUTION CODE Unclassified--Unlimited Subject Category 08 13. ABSTRACT (Mu;rnum 200 words) The Linear Aerospike SR-71 Experiment (LASRE) is presently being conducted to test a 20-percent-scale version of the Linear Aerospike rocket engine. This rocket engine has been chosen to power the X-33 Single Stage to Orbit Technology Demonstrator Vehicle. The rocket engine was integrated into a lifting body configuration and mounted to the upper surface of an SR-71 aircraft. This paper presents stability and control results and performance results from the envelope expansion flight tests of the LASRE configuration up to Mach 1.8 and compares the results with wind tunnel predictions. Longitudinal stability and elevator control effectiveness were well-predicted from wind tunnel tests. Zero-lift pitching moment was mispredicted transonically. Directional stability, dihedral stability, and rudder effectiveness were overpredicted. The SR-71 handling qualities were never significantly impacted as a result of the missed predictions. Performance results confirmed the large amount of wind-tunnel-predicted transonic drag for the LASRE configuration. This drag increase made the performance of the vehicle so poor that acceleration through transonic Mach numbers could not be achieved on a hot day without depleting the available fuel.
14. SUBJECT TERMS 15. NUMBER OF PAGES Aerospike rocket, Flight test, LASRE, SR-71, Stability and control 16. PRICE CODE A03 17. SECURITY CLASSIFICATION I 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified Unlimited NSN 7540-01-280-5500 l Standard Form 298 (Ray. 2-89) Prescribed by ANSI Std. Z39_18 298-102