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NASA Technic a l M emor a ndum 58223 NASA-TM-58223 1980001687 6
i i o " Orbiter / ShuttleC a rrierAircraftSepar a uon:
WindTunnel, Simulation,and FlightTest
Overview and Results
D. J. Hom a n, D. E . Denison,
and K. C . Elchert _O_ i _E, F E p _E / _C E
May 1 9 80
J
] : J tJ L 8 1980
" L ANGLE Y RESEA R CH CE N T Et _ N _A LIBRARY , NASA National Aeronautics and Space Administration
LyndonB . JohnsonSpaceCenter
Houston , Texas77058
NASA Technic a l M emor a ndum 58223
., Orbiter / Shuttle Carrier Aircr a ft Separation:
Wind Tunnel, Simulation, and Flight Test
Overview and Results
D . J . Hom a n
L yndon B. Johnson Space Center
Houston, Texas
D. E. Denison a nd K. C. Elchert
Rockwell International SpaceDivision
Downey, California
NI A
National Aeronautics and SpaceAdministration Scientific a nd Technical Information Office CONTENTS S e c ti o n P age FREE FLIGHTS ............................ I0 Tail cone-On Fli g hts ........................ II
C ON C L UDI N G R EMA RK S 15
iii TA B LE S Table Page I WIND TUNNEL TESTS ..................... . :17 III ORBITER TAIL C ONE-ON S EPARATIONPARAMETERS .......... 18 " %, iv FIGURES Figure Page 1 Mated Orbiter / SCA configuration • (b) Basic vehicle dimensions and configuration 7 Inert-captive flight 59 separation data run I (d) Orbiter pitching moment coefficient_ CMo , time (g) SCA pitching moment coefficient , CMc _ time 8 Elevon bias effect on Orbiter pitch acceleration and 9 SCA coefficients compared to Orbiter elevon deflection . . . 32 II Separation initial conditions from captive-active v Figure Page 12 Free flight I , tailcone o n o (a) Orbiter pitch rate command, @cmd_ time history ..... 3 5 13 Free flight 2 , tailcone on 14 Free flight 3, tailcone on vi Figure Page 16 Free flight 3 SCA / Orbiter tailcone-on separation trajectory eeoeeeeeeeeeeeeeele (2 frames per second') 48 ! 7 Free flight 4 G O / N O - G O sep a ration window (a) Relative normal load factor compared to Orbiter pitch (b) Relative normal load factor compared to relative 1 8 Free flight 4, t a ilc o ne off (d) Orbiter angle of attack_ _o_ ti m e history ....... 5 1 (f) SCA normal load fa c tor, NZc , time history ....... 5 2 (g) Relative normal load factor, ANz, time history ..... 5 2 (h) Orbiter aft attach point separation trajectory ..... 5 3 19 Free flight 5, tailcone off (a) Orbiter pitch rate co m mand, 8cmd, time history ..... 5 4 (b) Orbiter pitch rate, 6 0 , time history .......... 5 4 (f) SCA normal load factor, NZc 9 time history ....... 5 6 vii Figure Page 20 Orbiter tailcone-off separation initial conditions (a) Relative normal acceleration compared to Orbiter pitch (b) Relative normal acceleration compared to relative 21 Free flight 59 SCA / Orbiter tailcone-off separation viii i S UMMARY The approach and landing test phase of the Space Shuttle Program pre- sented the problem of carrying the Orbiter aloft atop the Shuttle Carrier _ Aircraft , a modified Boeing 747 , and separating the two vehicles in a safe and reliable manner. Five Orbiter free flights were flown using basically the same separation procedures. These separation procedures were designed using analytical prediction techniques and mathematical modeling that were the result of 3 years of scale-model wind tunnel testing , engineering anal- ysis , engineering simulations , and vehicle flight testing. The wind tunnel testing provided the initial information for building a separation aerodynamic data base to support both off-line digital simulations and man-in-the-loop simulations. The simulations served as a starting point for separation pro- cedure formulation and crew training. Verification of the aerodynamic data base and mafhematical modeling techniques was accomplished with flight test data retrieved th=ough the Shuttle Carrier Aircraft load measurement system.
The load measurement system also provided data for real-time assessment of separation parameters during the actual flights. Comparison of actual sepa- ration trajectories and analytically predicted trajectories revealed excellent agreement between the two and instilled confidence in testing and prediction techniques to be used to support the orbital flight test phase of the program.
INTRODUCTION During the approach and landing tests conducted at the NASA Dryden Flight Research Center (DFRC) 9 Edwards , California 9 the Orbiter was required to sepa- rate unpowered from the Shuttle Carrier Aircraft (SCA) , relying only on the lift generated by its wings . The mere size of the Orbiter (figs. l(a) and l(b)) created the problem of finding an aircraft capable of carrying it to a suitable altitude for release. This problem was solved by modifying a Boeing 747 specifically for carrying and launching the Orbiter. The solution to the problem of separating the unpowered Orbiter from the SCA is the subject of this report.
This paper discusses (I) the wind tunnel testing and engineering simula- tions used to generate the aerodynamic data base and flight test procedures , (2) the load measurement system (LMS) and flight test program used to generate flight data for comparison and verification of predictions_ and (3) the re- suits of the Orblter / SCA separations performed during the approach and landing test (ALT) program. The appendix is an analytic discussion of the computer program.
It should be noted that Alan L. Carter , DFRC , was responsible not only for the successful incorporation of the LMS into the flight test program but also for the acquisition of the SCA data required by the NASA Lyndon B.
Johnson Space Center (JSC) and Rockwell International Space Division engi- neers during and following each flight. The calibration data and equations for the load measurement system , along with all postflight corrections to the data, were available only through his personal efforts.
George M. Glenn, McDonnell Douglas Corporation, Houston Astronautics Division , Houston , Texas, assisted in the development of the postflight data reduction program and was instrumental in the overall success of the separa- tion aerodynamics subsystem for the ALT program.
In compliance with NASA's publication policy , the original units of measure have been converted to the equivalent value in the Systeme Interna- tional d'Unites (SI). As an aid to the reader, the SI units are written first and the original units are written parenthetically thereafter.
SYMB O L S a acceleration , m / sec 2 c mean aerodynamic chord , m CD drag c o effi c ient, D / q S C L lift coeffi c ient, L / q S CM pitching moment coefficient, M / qS_ LBref Orbiter body reference length , m N Z normal acceleration, g linear acceleration, m / sec 2 T time , sec V velocity , m / sec (knots) equivalent airspeed angle of attack , deg AN X relative longitudinal acceleration, g AN Z relative normal acceleration , g AX relative longitudinal separation distance , m AZ relative vertical separation distance, m _e Orbiter elevon deflection, deg @ pitch attitude , deg pitch rate9 deg / sec pitch a cceler a tion, deg / sec 2 Subscripts: o Orbiter c Shuttle C arrier Aircraft c md co m mand . s ep separation OVERVIEW Figure 1 shows the Orbiter mated to the Boeing 747 that was modified for use during the ALT progr a m and give s basic vehicle dimen s ions and configura- tion details. Figure 2 is a flow diagram that traces the testing and analysis aspects of the s eparation problem that is discussed here in terms of wind tunnel testing, simulations , the load measurement system , inert flights , captive-active flights , and free flights.
Approximately 1400 hours of wind tunnel testing provided the aerodynamic data base from which analyses and off-line simulations were used to generate the first procedures required for separation. This initial look at procedures determined the data requirements and planning for the c aptive phase of the flight test program. Man-in-the-loop simulations incorporated the results from previous an a lyses with the experience of the respective flightcrews to (I) optimi z e the separation procedure s and te c hniques , ( 2 ) gener a te simulated flight data to debug and exercise postflight data reduction programs, and (3) give some insight into the implementation of the information to be gained during the actual flight program. Four of the eight captive flights produ c ed useful separation data that were incorporated into previous analyses to obtain the final set of separation condition s and procedures to be used on the five ALT flights. The final section of the report contains a discussion of the free-flight test results.
WIND TUNNELTE S TING w The Orbiter / SCA data base was developed through an abbreviated schedule of wind tunnel test s performed during a 22-month period. The final Separa- tion Aerodynamic Data Book was published in November 19 7 6 (ref. i). The te s ts were divided into four major categories as shown in table I.
The first category , configuration development , involved a s et of tests designed to gather data on various proposed configurations as a first approx- imation at optimizing both mated vehicle climb performance and Orbiter / SCA separation performance. Two basic Orbiter configurations (with and without the tailcone I) were tested; a range of Orbiter incidence angles and Orbiter elevon deflections was covered; and various drag-reducing attach structure fairings were assesse d . The tests were c o nducte d using two model scales , two facilities , and a range of Mach numbers and Reynolds numbers to provide a means for correlating and abbrevi a ting future tests throughout the program.
The '_ , ated data base" test provided performance , stability , and control data for the mated vehicles in the launch configuration. The test also pro- vided basic isolated SCA data and proximity data for both vehicles at a sep- aration distance of zero meters , or at the instant of separation. The tailcone-on Orbiter configuration was used f o r the entire test. Orbiter elevon and body flap effects and S C A stabilizer effects on the proximity data were obtained during this test. Data from this test were used as a basis for establishing initial target conditions for separation.
The 'bated verification" test used the same model as the mated data base test but was performed in a different facility. This test replicated runs from the mated test to establish confidence in previous data and expanded the mated data base w ith the tailcone-off Orbiter data.
The "separation data base" test provided data for decaying the proximity effects of each vehicle on the other, from their maximum influence in the mated configuration to free stream where neither vehicle influen c ed the other.
Data were taken with the S C A and Orbiter mounted on separate balances and stings to allow the vehicles to be po s itioned at various distances apart in the same tunnel at the same time.
A matrix of the basi c c onfigurations tested during the program is shown in figure 3. The amount and quality of data obtained during these tests and the sensitivity of these data permitted elimination of two complete tests from the wind tunnel test program.
SIMULATIONS The separation aerodynamic data base developed through the wind tunnel test progr a m was incorporated into the off-line and man-in-the-loop engineer- ing simulations that were used to formulate the separation maneuver.
Two off-line digital simulations were implemented on computer systems at the Rockwell International Space Division and at JS C . The major emphasis of these programs was to perform parametric studies of the separation maneuver to determine how it was affected by airspeed, Orbiter elevon setting9 Orbiter incidence angle, Orbiter center-of-gravity (c.g.) location, the tailcone, aerodynamic data tolerances, proximity aerodynamics, winds, turbulence, '_ IThe tailc o ne was an aerodynamic fairing placed over the aft end of the Orbiter to reduce drag and S C A tail buffet and improve the performance of the m_ted vehicle.
_" 4 nonequilibrium conditions , and premature separation of one or more attach points. The studies were also used to evaluate emergency Orbiter jettison procedures during the inert and ferry flight tests. The off-line simulations incorporated mathematical models of each vehicle's flight control system9 equations of motion , and algorithms for combining the aerodynamic data needed to "fly" the simulations. The computer program that was run at Rockwell International Space Division was formulated and checked out at McDonnell Douglas Corporation_ St. Louis 9 whereas the program run at JSC was a modifi- cation of the existing space vehicle dynamics simulation (BVDS) program.
The SVDS program was modified (with Orbiter / SCA data_ mass properties , etc.)
by McDonnell Douglas Astronautics Company_ Houston (MDAC-H) , and both programs were compared for compatibility of results.
Five man-in-the-loop simulations were run to obtain engineering data on the best way to proceed with the mated and separation sections of the flight test program. These simulations introduced_ into the analysis and planning , the pilot techniques required to assure good separation of the two vehicles.
The first of the simulations was a two-body , six-degree-of-freedom , man- in-the-loop real-time flight simulation. The SCA crew station was a fixed- base simulator with an out-the-window display of the horizon9 whereas the Orbiter crew station was a moving-base fighter cockpit station that was mod- ified to include a rotational hand controller (RH¢) and dedicated separation switches and displays. The simulator had the capability of representing the flight characteristics of the SCA and the Orbiter during the mated flight 9 the separation transient , and the free flight of each vehicle.
The next three simulations were flown in fixed-base simulators. Two of the simulations used a manned Orbiter crew station with a "canned" SCA tra- jectory_ and the third used a manned $¢A crew station with a canned Orbiter trajectory. The canned portion of each simulation was computer generated using mathematical models of the flight control system and equations of motion similar to those used in the off-line simulations. The SCA simulation was flown by the SCA pilots and was used to investigate pilot capabilities for attaining separation initial conditions , optimum procedures to minimize altitude loss during the "pushover" maneuver , and procedures for acquiring separation data during the captive-inert and captive-active flight phases the of ALT.
The Orbiter simulations were run in the avionics development laboratory at Rockwell International Space Division , Downey , California , and in the crew procedures evaluation simulator at JSC. These simulations investigatedthe effects of pilot steering techniques_ aerodynamic variations , configuration variations , winds , gusts , and turbulence on the Orbiter separation.
The fifth simulation was flown at DFRC in conjunction with a wake vortex study that used the Boeing 747 before it was modified to carry the Orbiter.
The 747 was equipped with smoke generators on each wingtip to mark the vor- tices created by the wings. F-I04 and T-38 aircraft were flown by the Orbiter crews and DFRC test pilots in formation with the 747 to simulate a nominal separation maneuver , from SCA pushover through Orbiter separation.
The Orbiter was simulated by the smaller planes that flew in the same verti cal and horizontal position as the mated Orbiter during the pushover_ but it was laterally displaced about 60 meters to the right of the 747. When the SCA pilot called "launch ready_" the simulated Orbiter (F-f04 or T-38) per- formed the nominal separation steering and the SCA performed its postsepara- tion bank maneuver. The flights were designed to obtain qualitative data on postseparation clearances between the two vehicles and to obtain vortex avoi dance d is tances.
LOAD MEASUREMENT SYSTEM The LMS for the Orbiter / SCA was developed by the Boeing Company9 Aero- space Division 9 Kent , Washington. The LMS was designed to measure and record the attach forces between the two vehicles during the mated portion of each flight. An overall view of the LMS is shown in figure 4. The load cells were thin-walled cylinders instrumented with strain gages to measure axial and shear forces 9 and were located on each of the three Orbiter / SCA attach struts (fig. 5).
Figure 6 shows an exploded View of the_ aft load cell. The forward load cell measured the relative vertical and side forces at the forward attach point. The left aft load cell measured the relative vertical and drag forces at the left aft attach point , whereas the right aft load cell measured the relative vertical_ drag_ and side forces at the right aft attach point. These forces were recorded onboard the SCA and also telemetered to the DFRC control room_ where they were displayed in real time on strip-chart recorders. The forces were also combined mathematically to display the relative normal (AN Z) and axial (ANX) accelerations between the Orbiter and the SCA and the instan- taneous Orbiter pitch acceleration (0). These strip-chart data provided quick-look information for rapid postflight analysis and provided a basis for making a real-time decision to separate on the initial tailcone-off free flight.
The recorded forces were used in conjunction with recorded time histories of the SCA attitudes , rates , and accelerations as input to a computer program.
The program was designed to take this information and calculate (for specified portions of each flight) the aerodynamic coefficients for the Orbiter in prox- imity to the SCA , the SCA in proximity to the Orbiter , and the mated vehicles.
The Ground Reduced Aerodynamic Coefficients and Instrumentation Errors (GRACIE) program was used to verify and adjust the aerodynamic data base pro- vided by the early wind tunnel testing with actual flight data. (See the appendix for a description of the GRACIE program.)
INERT FLIGHTS The inert-captive phase of the ALT program was flown with two separation objectives: first , to demonstrate the Orbiter / SCA airworthiness within the operational envelope required to accomplish the ALT and , second , to conduct a preliminary evaluation of the Orbiter launch profile and procedures. Three taxi tests and five flights were performed with the Orbiter unmanned and un- powered and with all control surfaces locked in position. The Orbiter was configured as follows for all five flights.
Weight 638 764 newtons c.g. 64.5 percent LBref Tailcone On Incidence angle 6° Elevon -I ° (up) Body flap -11.7 ° Rudder 0° Speed brake 0° The taxi tests were performed with the mated Orbiter / SCA configuration to evaluate handling qualities during the takeoff roll , and braking and steering performance during the landing roll.
Following these tests , the mated configuration was flown five times.
The first four inert flights were flown to obtain takeoff and climb perform- ance data; to investigate stability and control envelopes, flutter response, and buffet and loads boundaries; and to perform airspeed calibration checks.
Inert flight 4 focused on evaluating configuration variables associated with the launch maneuver 9 as reflected by the buffet levels and aircraft handling characteristics. During this flight, the SCA inflight spoilers were deployed for the first time and the aircraft performance was assessed based on the special thrust ratings on the engines. These two items were of major importance to the separation maneuver because the special rated thrust (SRT) increased the climb ceiling for the mated configuration , allowing separation to occur at a higher altitude, and the inflight spoilers decreased the lift on the SCA just prior to separation, creating a high relative normal accelera- tion between the two vehicles at separation. This flight provided engineers " their first look at a separation-related parameter in the form of the incre- mental effect of the inflight spoilers on each vehicle in close proximity.
" Inert flight 5 obtained data during two simulated launch maneuvers starting at ceiling altitude and terminating after approximately 20 seconds of steady-state data following the "launch ready" call by the SCA pilot.
Both vehicles were configured as they would be for an actual separation with the exception of the Orbiter elevon , which was locked at -I ° (up). This elevon position was chosen as the optimum position for jettisoning the un- manned Orbiter in the event of an emergency and for providing flight data with the Orbiter elevon close to the predicted position for separation. The emergency jettison capability was never confirmed or required.
During the launch maneuvers_ "launch ready" was called when the SCA had reached equilibrium glide conditions with the inflight spoilers deployed_ the engines in idle thrust_ and the airspeed at 139 m / sec (270 knots) equivalent airspeed (EAS). Acceptable launch conditions were actually considered to be a velocity at separation of 139 +_ 2.6 m / see (270 ± 5 knots) EAS and a normal acceleration of ig ± 0.3g. The SCA pilot was able to control the mated vehi- cle well within these constraints throughout the entire data acquisition period on both launch attempts.
• Data obtained during this flight (using the LMS and GRACIE) are shown in figures 7(a) to 7(g). Based on these flight data_ it was discovered that there was some confusion about the correct SCA data base. The problem was traced to incorrect use of wind tunnel incremental data_ and the separation aerodynamic data were updated to reflect the actual flight data.
The inert-captive flight program was accomplished with a total flight time of approximately II hours 36 minutes9 and all flight test requirements were satisfied within the flight envelope tested. The simulated launch maneu- vers on inert flight 5 verified that (I) the Orbiter ! SCA configuration could achieve and stabilize on the separation parameters using the prescribed pro- cedures without exceeding Orbiter or SCA constraints , (2) safe separation initial conditions could be achieved with the baseline separation configura- tion and a irspeed_ and (3) the mated configuration could recover from an aborted separation maneuver within the vehicle constraints.
CAPTI VE-ACTIVE FLIGHTS Three captive-active flights were flown with the Orbiter manned. The objectives of these flights were to verify (I) the separation configuration and procedures; (2) the integrated structure_ aerodyDmmics_ and flight con- trol system; and (3) the Orbiter integrated system operations.
The first flight in this series_ designated CA-IA_ was limited to an airspeed o f 93 m / sec (180 knots), which required the SCA flaps to be lowered throughout the entire flight and thus precluded the acquisition of any useful LMS data for separation analysis. With the flaps in any position other than completely retracted_ the SCA developed enough added lift to overcome the lift generated by the Orbiter_ resulting in the att a ch struts and load cells always remaining in compression. The speed restriction and flap position also created a problem in jettisoning the Orbiter in the event of an emer- gency. Therefore_ the procedures for emergency separation required that the flaps be retracted and the airspeed be greater than 113 m / sec (220 knots) EAS with the SCA inflight spoilers deployed. Based on studies using the off-line simulations_ emergency separation procedures were devised to allow a safe sep- aration for both vehicles. The procedure called for the SCA pilot to acceler- ate while retracting the flaps. Once the flaps were retracted and the air- speed was in excess of 113 m / sec (220 knots) EAS, the engines were to be idled and the s poil e rs d e ployed , f ollowe d by a "launch ready" call from the SCA pilot. At that time , the Orbiter crew could command separation and use the nominal steering procedures. If emergency separation was time critical , the procedure would have been for the SCA pilot to retract the flaps , deploy the spoilers, and pitch over to NZmated _< 0.3g; then the Orbiter crew would be cleared to separate.
The second captive-active flight , CA-I, expanded the flight envelope for both vehicles and provided a range of Orbiter elevon position data in the launch configuration. After clearing the vehicles for flutter and buffet through an airspeed of 139 m / see (270 knots) EAS, the mated pair performed a separation run to gather data for analysis of separation configurations for the upcoming tailcone-on free-flight tests. At approximately 40 minutes into the flight , the Orbiter / SCA configured for the separation data run. The SCA started a gradual pushover to accelerate to a 139 m / sec (270 knots) EAS equi- librium target condition. When the airspeed increased to approximately 2.6 m / see (5 knots) less than the target speed, the SCA pilot put the engines in idle thrust and deployed the inflight spoilers to achieve the equilibrium glide conditions required for the data run. The Orbiter was configured as follows.
Weight 667 233 newtons c.g. 63.8 percent LBref Taflcone On Incidence angle 6° Body flap -9.7° Rudder 0° Speed brake 5° The Orbiter elevon was trimmed to 0° for the first data point and held for 5 seconds. The RH¢ was then moved full forward and held for approximately 5 seconds of steady-state data. (Elevon software limits during this portion of the flight restricted the elevon travel between-1.5 ° (up) and +1.5 ° (down) only; therefore , moving the RHC full forward drove the elevon to the - +1.5 ° (down) position.) In a similar manner, the elevon was positioned at -1.5 ° by moving the P, HC full aft and holding it for approximately 5 seconds of data. The R H C was then returned to the detent position (0°) followed by a full right movement to put in I° of right aileron for approximately i0 sec- onds , after which the Orbiter commander terminated the separation data run.
The SCA pilot then performed a gradual recovery and reconfigured for landing.
Data retrieved from the flight, using the LMS and the GRACIE program, are presented in figures 8 , 99 and I0. The primary separation parameters , relative normal load factor (ANZ) and Orbiter pitch acceleration (@o, are shown in figure 8 for the elevon positions tested. These data and the aero- dynamic coefficient data presented in figures 9 and i0 indicate a shift between the predicted values and the flight test data. The shift seemed to be equivalent to an approximately -I ° bias in the Orbiter elevon position; i.e. , the data indicated that instead of being at 0° , the elevon was actually at -I °. The data in figure 9 also show that the elevon effectiveness was in excellent agreement with the preflight predictions. The size of the separa- tion window for the first two tailcone-on free flights was large enough to handle the -I ° elevon bias without affecting the separation drastically. It was expected that data from the remaining captive-active flight and the first two free flights would give some insight as to the cause of the shift in data.
The elevon effectiveness results were used to determine the correct ele- yon setting for separation on the third free flight_ which was flown with a more aft c.g. location and therefore required more down elevon.
The final captive-active flight , ¢A-3_ was a dress rehearsal for free flight 1 up to the point of separation. The pushover was performed as it would be on free flight 1 with the SCA and Orbiter configured for launch.
The Orbiter crew performed the preseparation functions of moving the elevon from the climb position (-2°) to the separation position (0 °) and commanding a +2 deg / sec pitch rate with the RHC. At the "launch ready" call, the maneu- ver was aborted and the SCA recovered and reconfigured for landing.
• The LMS data and the GRACIE program were used to generate the separation parameters at the time of "launch ready" for postflight analysis. A compari- son of CA-I and CA-3 data_ shown in figure ii , indicates the elevon bias was not apparent on CA-3. This finding gave rise to questfons regarding data repeatability and el evon position calibration accuracy. No changes to the proposed separation configuration were made because of the relative insensi- tivity to small elevon dispersions in the first two free-flight separation profiles. Also 9 two more repeat data points would result on free flights I and 2 , from just after the "launch ready" call to the instant of separation.
All separation-related data acquired during the captive-inert and the captive-active flight test phases indicated that the desired conditions for separation would be attained by the baseline vehicle configuration and that separating the Orbiter frOm the SCA safely would not be a problem.
FREE FLIGHTS The primary objectives of the free-flight phase of the ALT's were to ver- ify (I) the handling qualities of the Orbiter vehicle, (2) the performance of the Orbiter subsystems , and (3) the Orbiter / SCA separation. The Orbiter / SCA separation is emphasized in this section. The free-flight Orbiter configura- tions tested are listed in table II.
Tailcone-On Flights On the first free flight of the Orbiter "Enterprise , " the flight pro- ceeded as expected up to the time of separation. The Orbiter / SCA pushover was initiated at an altitude of 8733 meters , followed by separation at 7644 meters. All went nominally until physical separation occurred , at which time general-purpose computer _mber 2 (GPC 2) failed. The SCA pilot called "launch ready" at an airspeed of 138 m / sec (268.3 knots) EAS and a pitch attitude of -6.4o; within I second, the Orbiter commander commanded separation. At separation , t_e initial relative normal acceleration of the two vehicles was 0.99g and the Orbiter pitch acceleration was 3.1 deg / sec2o The nominal steering command for the Orbiter was to have been as follows: w I. Comm a nd a pitch rate of +2 deg / sec for 3 seconds.
2. Command 0 deg / sec for 2 seconds.
3. Bank right 20° .
4. Push over at -I deg / sec.
Figure 12(a) shows that the actual commanded pitch rate was initially about 2.5 deg / sec but peaked to about 5 deg / sec at I second during the transient following separation. This command resulted in a pitch rate , as shown in figure 12(b), of approximately 5 deg / sec at 1.3 seconds and a higher-than- nominal angle of attack, II.I ° maximum compared to 9.3 ° nominal (fig. 12(d)).
The Orbiter normal load factor (fig. 12(e)) peaked at 2.1g 1.7 seconds after separation and thus violated the Orbiter constraint of N Z < 2.0g. The rea- son for the high initial pitch rate command is unknown; however, the initial transient due to the rapid change in the normal and pitch accelerations and the master alarm triggered by the loss of GP¢ 2 were more than likely the major contributing factors. Also , the ALL SEP B indication and the backup separation discrete were not seen by the flight control system because of the loss of GPC 2 ; these are the signals from the separation switches on the aft attach points and the pilot's keyboard , respectively , that enable the primary flight control system (PFCS) normal control surface limits. Follow- ing free flight I, the ALL SEP B and the backup separation discrete were no longer handled by the same GPC.
Postflight analysis of LMS data indicated that, as on CA-3, the elevon bias was not apparent. Reconstruction of the free flight 1 separation trajec- tory was performed using the off-line simulation programs with the vehicle initial conditions and the Orbiter steering command as inputs. (See fig- ures 12(a) to 12(g).) The initial conditions were obtained from LMS and GRACIE data and from 8CA flight instrumentation; the Orbiter pitch rate com- mand was obtained from Orbiter downlist data. Also shown in figures 12(a) to 12(g) is a comparison of flight and predicted separation parameters. The predictions are based on the postflight data of pitch rate command. The off- line data show excellent agreement with the flight data. The difference seen in the SCA normal load factor (fig. 12(f)) is attributable to the difference between the postseparation steering maneuver used by the SCA pilots and that programed into the off-line simulation. Figure 12(h) is the Orbiter aft ii attach point trajectory time history as reconstructed by the off-line simu- lation programs based on the postflight data. Photographic data obtained on free flight 1 were not sufficient for comparison with the predicted 'tra- jectory. On all subsequent flights_ adequate photographic data were available.
Free flight 2 was flown with the Orbiter c o nfigured as in free fligh t I.
The flight itself was designed to include a number of aerosurface inputs_ programed tes t inputs , and other data-gathering tests , but the separation procedure and configuration were identical to that of free flight Io The Orbiter / SCA pushover maneuver was initiated at an altitude of 8887 meters9 with separation occurring at 7718 meters. The Orbiter separated at an air- speed of 138.5 m / sec (269.2 knots) EAS with an initial relative normal load factor of 0.96g and a pitch acceleration of 2.4 deg / sec 2. The actual pitch rate steering co m mand is shown in figure 13(a); figures 13(b) to 13(g) com- pare the values generated off-line9 based on postflight data from the LMS and GRACIE. The predicted data again display excellent agreement with the flight data and 9 as seen in figure 13(h)_ the predicted Orbiter trajectory agrees with the flight photographic data. The flight photographs were taken with a motor-driven 35-millimeter camera at a speed of 2 frames per second.
The first frame Of the separation sequence was not necessarily taken at the instant of separation; therefore_ the first point on figure 13(h) represents a time somewhere between. Tse__ and Tsep + 0.5 second. Subsequent points occur at 0.5-second intervals.
Free flight 3 was the third and final Orbiter free flight to be flown with the tailcone covering the aft end. This Orbiter configuration was changed by reballasting to obtainla more aft c.g. location_ 65.9 percent reference length. The incidence angle remained at 6° , but the launch air- speed and the elevon position for separation were changed to acco m modate the aft e.g. location. The Orbiter elevon had to be lowered from the free flight 1 and free flight 2 position to counteract the increased nose-up pitching moment caused by the more aft c.g. location. The elevon deflection originally was to have been +1.5 ° , but based on the apparent elevon bias noted on CA-19 the separation setting was loaded into the flight computers as +2.5 °. Also , following free flight i , the separation airspeed was lowered from 138 m / sec (268 knots) EAS to 129 m / sec (250 kno t s) EAS because the lower speed would provide safe separation conditions and reduce the possibility of violating the Orbiter normal load factor constraint of NZ < 2.0g as on free flight 1. When data from CA-3 and free flights i and 2 failed to indicate the elevon bias noted on CA-19 the elevon setting for free flight 3 was re, evaluated and it was concluded that , with the lower launch speed , the elevon position of +2.5 o would provide acceptable sep a ration conditions and would not warrant reloading the flight computers with the original elevon setting.
The Orbiter / SeA pushed over at 8689 meters and the Orbiter was released at a pressure altitude of 7937 meters. The higher launch altitude was a di- rect result of the reduction in separation airspeed. The airspeed at the "launch ready" call w a s 129.9 m / sec (252.7 knots) EAS with a relative normal load factor of O.92g and an initial Orbiter pitch acceleration of 1.0 deg / sec2; these data were well within the separ a tion window targeted by the SCA pilot. Subsequent postflight analysis and trajectory reconstruction resulted in the data presented in figures 14(a) to 14(h).
The separation results from the tailcone-on free flights were in excel- lent agreement with the off-line predictions and well within the constraints of the separation window 9 as shown in figure 15. Table III summarizes the pertinent separation parameters with a comparison to both the target values and the predicted values. The predicted values are based on the actual condi- tions at separation , and the target values are the conditions to which the SCA pilot attempted to fly the mated vehicle.
Figure 16 shows a typical tailcone-on separation trajectory sequence as viewed from the SCA chase plane. The sequence was shot at 2 frames per second.
Tailcone-O ff Flights The last two flights of the ALT program were flown with the Orbiter con- figured as it would be when returning from orbit; i.e._ with the tailcone removed and with the three main engine bells in place. The removal of the tailcone presented two major problems with the separation phase of the flights. First , without the tailcone9 the increase in the buffet level could possibly result in an SCA cockpit environment that would make it impossible for the SCA pilot to attain the specified target conditions. Second 9 with the removal of the tailcone , the change in Orbiter pitching moment required +7 ° of down elevon , which was well outside the elevon range tested in the preceding flights. The SCA tail loads and climb performance degradation created by the increased buffet and drag levels_ respectively_ were also un- knowns that could have terminated the flights prior to separation. A fourth captive-active flight was originally planned to investigate the flight enve- lope of the tailcone-off configuration but was deleted. The objectives of the canceled captive-active flight were combined with free flight 4 and were eval- uated in the first half of the flight. Also , the Orbiter incidence angle was left at 6° instead of changing it to 5° as originally called for in the pro- gram. This created one less unknown to be verified for the tailcone-off flights and also relieved some of the concerns of buffet by lowering the tar- get launch speed from 139 to 126 m / sec (270 to 245 knots) EAS.
The first portion of free flight 4 was dedicated to a real-time assess- ment of the buffet-induced loads and verification of the separation configura- tion and target conditions. A real-time GO / NO-GO decision for separation was made based on LMS data telemetered to the ground and displayed on strip-charts in the DFRC control room. The LMS data were also filtered to reduce the noise in the strip-chart parameters caused by the increased buffet levels.
The increased drag of the tailcone-off Orbiter introduced a third separa- tion parameter , the relative axial load factor 9 which had been insignificant during the initial portion of the tailcone-on Orbiter separation. Two inde- pendent studies , based on the Rockwell International Space Division and the JSC off-line simulations , were performed to investigate possible combinations of the relative normal and axial load factors and Orbiter pitch accelerations that would allow a safe separation. The resulting separation window is shown in figures 17(a) and 17(b).
A step-by-step analysis of the buffet-induced loads and vibrations from take-off through the maximum airspeed of 129 m / sec (250 knots) EAS led to the initiation of the separation data run 9 which provided the information neces- sary for the separation GO / NO-GO decision. With the Orbiter elevon positioned at +7o9 the SCA pilot deployed the inflight spoilers , idled the thrust , and attained equilibrium glide conditions at a n airspeed of 127 m / sec (247 knots) EAS. Approximately 10 seconds of steady-state data w a s taken before the SCA and Orbiter reconfigured and beg a n their climb to the pushover altitude for separation° During the interim , the strip-chart data were evaluated and it was concluded that the separation configuration provided adequate condi- tions to assure a safe separation. The "quick-look" results are shown in figures 17(a) and 17(b). (The load cells contained two sets of vernier load measurements for separation_ which provided redundant readings for each param- eter.) If the data had not fallen within the acceptable regions of the sepa- ration windows9 the SCA / Orbiter would have performed a second separation data run to obtain data at vari ou s elevon positions , as was done on CA-I 9 so that the optimum setting for separation on a subsequent flight could be chosen.
Having received the "go ahead" for separation , the SCA initiated the pushover maneuver at 7397 meters , with separation occurring at approximately 6541 meters at a velocity of 127.4 m / sec (247.7 knots) EAS. The relative normal and axial accelerations were 1.04_ and 0.27g , 2 respectively , and the Orbiter pitch acceleration was 0 deg / sec z. The comparison of the flight data and the postflight off-line analysis (figs. 18(a) to 18(g))shows excellent agreement, as does the separation trajectory data (fig. 18(h)).
The final flight in the ALT program_ free flight 59 was flown with the same Orbiter configuration as in free flight 4. The prime objective of free flight 5 was to land the Orbiter on the concrete runway, which required sep- aration to occur at a predetermined point in the sky. The separation target conditions were the same as those of free flight 4. The SCA initiated push- over at approximately 6632 meters with the SCA pilot calling "launch ready" at approximately 6041 meters. The Orbiter crew co m manded separation approx- imately 7 .5 seconds later at 5791 meters. The airspeed at that time was 128.9 m / sec (250. 7 knots) EAS with a relative normal acceleration of 1.0g , a relative axial acceleration of 0.17g , and a pitch acceleration of -1.0 deg / sec 2.
The trajectory reconstruction using the actual initial conditions and vehicle configurations at separation again showed excellent agreement lamong flight data, off-line simulation data 9 and photographic data (figs. 19(a) to 19 (h)).
2There was a bias in the axial load channels from the LMS as seen in the data after separation , which , if accounted for , would decrease the calculated _X from 0.27g to 0.17g at the instant of separation.
A compilation of the separation results from the tailcone-off flights is presented in figures 20(a) and 20(b) and table IV. Figure 21 shows a typical tailcone-off separation sequence. Notice the more aft relative motion of the Orbiter without the tailcone as compared to the Orbiter motion with the tail- cone in figure 16.
CON C L U DING REMARKS The analytical prediction techniques and mathematical modeling incorpo- r a ted in the design of the separation procedures for the Orbiter / SCA were • b a sed on scale-model wind tunnel test data. These techniques proved to be e x tremely accurate and useful throughout the approach and landing test program.
The two-body man-in-the-loop si m ulations, with one vehicle trajectory c anned, provided adequate fidelity for crew training and c rew inputs to the sep a r a tion procedures and configur a tions.
The lo a d measurement system installed aboard the SC A provided a means for extr a cting the proximity aerodynamics and was a reliable sour c e for making real-time assessments of separation and load s parameters. The load me as urement system also allowed some wind tunnel tests to be deleted from the program, with a c tual flight data completing the aerodynamic data base. The Orbiter separated from the SCA, suc c essfully and as predi c ted_ five times during the approach and landing test program.
Lyndon B. Johnson Sp ac e C enter National Aeronautics and Space Administration Houston, Texas, April 2, 1980 953 - 36-0 0- 00 - 72 R EF ER ENC E I. O rbiter /7 4 7 Ca rrier S eparation A erodynamic D ata Bo o k - SD M Ba s eline. • S D 7 5- SH -0033 C 9 Rockwell International Space Division 9 N ov. 19 7 6.
TAB L E I. - WIND TUNN EL TESTS ( S e e f i gu re 2 .)
Test obj ec tive T e st n u m b er Location a Model scale Test date C onfig u ra t ion C A 5 BTWT 0.0 3 S ept. 1 9 7 4 deve lo p m e n t C A6 B TW T . 0 3 M a y 197 5 CA23A A R C • 0 1 25 Ma r ch 1 9 7 5 Ma t ed da t a base CA I4 A B TWT . 03 N o v . 19 7 5 Mated v e ri f ic a tion C A13 AR C .03 Jun e 1976 S e pa r a tion da t a bas e C A2 0 BTWT . 03 O c t. 1 9 7 4 C A 23 B AR C .01 25 July 1 9 7 5 CA26 LTV .0 1 25 Au g . ! 9 7 5 aL o ca tio nsa r e as follo ws: ARC - Ames R esea r ch Cen t er , M of fe t t Fi e l d, Ca lif . ; B TWT - Bo ei n g Tra n s o n i cWin d Tunne l, Sea t t l e , Wash. ; a n d LTV - LTV Ae ro space Co r por a tion, Da ll as 9 Tex.
TAB L E II. - FREE FLI GH T ORBITER C ONFI GU RATIONS Flig h t Tail c one Weigh t , c .g. lo c ati o n, Orbiter in c iden c e N per c en t referen c e a ngle, deg length i O n 66 7 411 6 3 .8 0 6 2 On 66 7 411 6 3 . 80 6 3 On 667 0 55 6 5 .90 6 4 0 f f 67 0 5 47 66.2 5 6 5 off 670 5 47 66. 25 6 TABL E T it . - ORBI TE R TAI L C O N E-O N SE PAR A T ION PARA M ETE R S Par a meter Free flight •I Free flight 2 Free flight 3 Orbiter c .g., per c ent ........... 63. 8 63 . 8 65.9 Orbiter elevon , deg ............ 0 0 2.5 SCA airspeed, m / sec (knots) EAS Flight ................. 138.0 (268.3) 138.5 (269.2) 1 2 9.9 (252.7) T arget .......... ....... 139 (270) 139 (270) 129 •( 2 50) SCA pitch attitude, deg co t ................. - 6.3 8 - 5.91 - 2.95 F_ Fligh Target ................. -6 -6 -3.5 SCA altitude (MSL) a, m .......... 7645 7718 7937 Relative normal loading factor , g Flight ............. .... 0.991 0.956 0.917 Predicted ................ 0. 994 0. 907 0. 857 T arget ................. 0.9 0.9 0.88 o Orbiter pitch acceleration_ deg / sec 2 Flight ................. 3.1 2.4 1.0 Predicted ................ 2.4 2.4 0.1 Target ................. 2.5 2.5 0.6 aMean sea level.
TA B LEIV.-OR BI TERTAILCO N E-O F F S E PARAT I O N PARAMETERS
Parameter Flight _4 Flight 5 Orbiter c, g., percent ........... 66.25 66.25 Orbiter elevon9 deg ............ 7 7 SCA airspeed, m / see (knots) EAS Flight ................. 127.4 (247.7) 128.9 (250.7) 126 (245) 126 (245) Target .................
SCA pitch attitude , deg Flight -5 2 5 - 6 07 Target ................. -6 -6 u D SCA altitude (MSL), m ........... 6541 5 791 Relative normal loading factor9 g Flight ................. 1.04 1.0 Predicted ................ I.23 I. I Target ................. i.0 I.0 Relative axial load factor9 g Flight ................. 0.17 0.17 Predicted ................. 0.19 0. 2 Target ................. 0.2 0.2 Orbiter pit c h acceleration9 deg / se c 2 • • . • • • • . • e • • • • • • • • • Flight 0 0 - i 0 Predicted -0 5 -0 5 Target ................. 0.6 0.6 ( a ) Photograph.
Figure 1.- Mated Orbiter/SCA c o n f i g u r a t i o n .
S C A O r bite r Measu r ement W ing V e r ti c a l Ho riz o nta l W ing Verti c a l A r ea , m2 511 77.1 136 . 6 24 9 .9 38.4 Span , m 59.6 9 . 8 21 . 9 23 . 8 8 Aspect ratio 6 . 96 1.25 3.60 2.265 1.675 Tap e rratio 0 . 356 0 . 340 0 . 250 0 . 200 0 . 404 - Sweep , deg 37 . 5 (1 / 4 _) 45.0 (1 / 4 _) 37.5 (1 / 4 _ , ) a45 a45 Dihedral , deg 7.0 7 . 0 b3.5 - I ncidence , deg 2.0 - + 5 to - 10 0.5 -_
MAC Cm 8 3 8 5 6 9 12 1 5 1
, • . .
a L eading e d ge .
CMean aerodynamicchord .
_ , --- Orbiter bT r ai l ing edge. __T ' , 70 . 7 -----23 . 8 6° /, 42 . 2 "1
!
21.8 and load ceils and load se l ls (b) Basic vehicle dimensions and configuration details.
Figure i.- Concluded.
• • Ca p ti v e fli g hts
Q W ind t unnel data "_
Anal y sis / si m ula t ions
Tol e ra nc esbas e d @ P o s tf li g h t d a ta
r ed u ct i on / a na l y sis
- o n w i nd t u nne l d at a
i ct ed s e pa r a t ion w indow •
m ., , /l,t _ . _
_ __ _ • O pen se p ara t i onw i n do w•
I conditions] - o n fl i g h t dat a
Tolerances based
Q Data required f r om
captive flights let
, - condi t ion s
Man-in-the-loop si m ulations i a I J
@ E xercise d ata / k _ • Criteria fo r mo difying
re d uction p_ _ separation configuration / conditions
Figur e 2 . - Separati o n analy s i s flow c hart.
ii 1974 1975 1976 Test , , I , , = , ' i ' , ' ' , 0 N D c onfigu r at i on J F M A M J J A S 0 N D i|im l M L, l i mI |m|_Ji I Ke|[]i U J F M A M J J A S i 11 1, 19 6 8 5 20 I " 7 12 1 1 8 12 ca rri e r Orb i te r (Tai Icone on) Mated ____ I_ _ _ !-_ ,m (Tai I c oneoff) Mated _ [' 6 _ _ I -_ [_
Separat,on _
Figure 3. - Wind tunnel configuration matrix.
l oad
s e nsi ngu nit l o a d
( r ed und a nt se n sin guni t
Pre set fo r war d me a s ur ments ) (re dun da n t
measure m ents)
L o ad meas ur e m ent
s y ste m Flight test c onsoles
s i gnal c o ndit io ner and electr i cal racks
and electrical cabling
Figure 4.- Orbiter / SCA load measurement system.
Load sensing unit
Telescoping tube
Adjusting
AdJustment bolt (each side) brace (each side)
Boeing 747 forward support struts
Load sensing unit
Side load strut
D r ag str u t
Side load strut
Verti c al str_ /
-. Drag strut __ //
Boe i ng 747 aft support struts
Figure 5.- Attach strut load c e ll locations.
- - : Load r i ng
St rain g age s
!
",._ I
Lo wer p l a te
N u t pla t e
!
0 i Figure 6. Aft load c ell.
I0-
' - - -___._o;_t ; r............... - --- ....
8 -
_ 6 -
u _ J
o 4 -
- :m SCA -
<
0 I I I I I
8:45:5i 52 53 54 55 56
Time , sec (P.d.t.i
(a) Angle o f att ac k, e, time history.
0.5 - o
'_ . 3
oN o
o Flight
.2 -
'- Pred i cted
i_
-- -- U n ce r tainty in flight data
o
O - I I I I I
8 : 4 5 :51 52 53 54 55 56
T i me, sec (P . d.t.)
(b) Orbiter lift c o efficient9 CLo _ time hist o ry.
Figure 7.- Inert cap t ive flight 5_ separation data run i.
2 7
F l i g h t
-- - Pre d ic t ed '
0.1- 0 - - - - U ncer t a i n ty i n i ' fl .igh t da t a
. 08
g , °m
= ,06 "
o
-_ . 0 4
0 I I 1 I I
8:45:5 1 5 2 53 54 55 56
Ti m e , s ec (P.d.t.)
(c ) O rbit e r dr a g coe ffi c i e nt, CDo , tim e hi s t o ry.
e 0 , I 0 - Flight
_' -- - P red i c t ed
., 2
=. , 08 .... U n ce rta i nty i n fl ig ht data
.__
o .06
aJ o
E .04
t" 0 m e .- o
- _ . 02 -
0 I I I I I
o 8: 4 5:51 52 53 5 4 55 56
Ti m e , s ec (P.d.t.)
( d) Orbiter pit c hing m o ment c oeffi c i e nt, CM o , tim e hi s t o ry.
Figure 7.- Continued.
2 8
0. 5 - -- -- Fl i gh t
--- Pre di c t ed
o .4 - Uncertainty i n f l i g ht d at a
L 3 e,- ° _ °_ u .2 ,_ < 0 I I I I I
8 : 45:51 52 53 54 55 56
Ti me, s e c (P.d.t.)
(e ) SCA lift co effi c i e nt, CL c , tim e h i s t o ry.
0 . I0 - - --- Fl igh t
-- -- Pred i c t ed (,. )
.08 - ----- U ncer t ain t yin f l ig h t da t a
(. 3 : e,-
o 0 6 - '
"- J_ "_ ' ,_r I,,I . ..
• -° "
_J o
o 04 -
.02 ""_- J _ " - - - V ' -
u' )
0 t .. I L__ I__ _d
" 8:45:51 52 53 5 4 5 5 5 6
Ti me, se c (P.d.t.)
(f) S C A dr a g coe fficient_ C D c _ time his t o ry.
Figure 7.- Continued.
------ Flight
.... P re dict e d
Q. 0 4 .... Un c ertaint y i n flight : data
C ,m , m • 1, 4 ,,,, i
0 2 B
q ,l L) e-
" 0 0
E o E ,
= -.0 2
e,-
N
, m
< -. 0 4 _ I I i, I
8: 4 5: 5 1 52 53 54 55 56
u _ T ime, se c (P.d.t.).
( g ) S C A p it c hin g moment coe ffi c i e nt_ C M c , time hist o ry• Figure 7.- Concluded•
C) Captive-active flight 1 data
C] Predicted values
I Uncertainty due to instrumentation errors
'o
\ . _"
_.2- 8 - \\ ____N z
1 . 0- 6-\ \i
N ::_ ¢ x j :=,
. . 8- ._. 4
• "_, \ \
r_
_o .6- oz 2- \ \
o -- \
\
.>- .4 - _ 0 - '
-= \" \
._ \ q \ )
.2 - _ -2 - \
o \ \
\
\ -
\
0- -4 - \
\
-6 I I
-2 0 2
Orbiter elevon deflection ,
, deg
e Figure 8.- Elevon bias effect on Orbiter pitch acceleration and relative normal load factor.
0 Captive-active flight 1 data
[] Predicted values
I Uncertainty due to instrumentation error
0.6 - 0.12 - O.IO -
0 I I O I '1 - . 02 0 I I
-2 0 2 -2 0 2 -2 0 2
Orbiterelevondeflection , Orbiterelevondeflection, Orbiterelevondeflection ,
8e , deg 8e , deg 8e , deg
Figure 9.- SCA coefficients compared to Orbiter elevon deflection.
(_) C a p ti v e-a c t iv e fl_g h t 1 data [ 3 Pre di c t ed va ] ue s l Uncer t ain ty due t o ins t rumen t a t ion errors Figure i0. - O rbit e r coefficients compared to elevon deflection.
12 '_ Sep a r at ion wi nd ow 7
_ 8
._. •
O "l "0 " CA1
: _ o 4 -
g
.o • 0 CA3 T a r get _ .
llJ
_ 0
e- __ ,
_ L 4
o - 8 I I I I I I
0 .2 .4 .6 .8 I .0 I.2
Relative normalloadfactor , A N z ,g F i Eu re ii.- S e para tio n -i n iti a l cond itio ns fr om c ap ti v e- ac tive f lig h t s.
5 c.g. = 63.8 percent
4 ae=O°
ZXNz= 0 . 991g
¢_° 8 ' = 3.085 deg / sec2
_ n 0 E
o
"1 -
I I I J
0 2 4 6 8
Time from separation, sec
(a) Or b iter p itch r a te c ommand_ e c md_ time h i s tory.
5 ---- Flig h t
4 --- P r e dic te d
¢) "o "o 1 . Q: )
-1
. -2 I I I I
0 2 4 6 8
Time f r om sepa r ation , sec
( b) Orbiter pitch rate, Co, time history.
Figure 12. - Free flight i, tailcone o n.
20 - . --.-- Flight
--- Predicted
10 - _ - "
_ , _ , 6 , , , • .
-10 I I I J
0 2 4 6 8
Time Fromseparation , sec
(c) Orbiter pitch attitude9 eo, time history.
12 ----- Flight
[- _ -- - P r ed i cted
8-
I I I I
0 2 4 6 8
Time from separation , sec
(d) Orbiter angle of attack, 0_0, time history.
Figur e 12.- Continued.
3 -- Flight --- P r edicted o N j Z I- I I I I 0 2 4 6 8 Time fromsepa r ation , sec (e) Orbiter normal load factor , NZo , time history.
2 - -- Flight --- Predicted o I - !
N I I I I 0 2 4 6 8 Timefromseparation, sec (f) SCA normal load factor, NZc , time history.
2 - -- Flight --- Predicted z <1
1_ I ,-. \
I I I I 0 2 4 6 8 Timefrom separation, sec (g) Relative normal load factor_ ANz, time history.
Figure 12.- Continued.
40 - ----- Predicted
2" 6 S e c
E
N 30
<3 3 . 2
g , o _m
_ 20
2 .4
o >
> 2.0 SCA vertical stabilizer
" 1 . 6
1.2
-0.8
-
-O.4 I
0 I0 20
R e l ative long i tudina l separat i ond i stance , Z_X , m
(h) Orbiter aft attach point separation trajectory.
(Origin is fixed at SCA aft attach point.)
Figure 12.- Concluded.
3 F c.g. = 63.8 percent
2 ae : O°
o ANz = 0.956g
o I 8 " = 2.433 deg / sec2
_- o _Dn o.j
-_ 0
_ " E o -1
-2 I I I I I
0 2 4 6 8 , I0
Time from separation, sec
(a) Orbiter pitch rate co_and, Ocmd_ time history.
3 -- Flight
2 --- Predicted
°6 -1 '_
-2 I _
0 2 4 6 8 I0
Time from separation, sec
(b) Orbiter pitch rate, _)o_ time history.
Figure 13.- Free flight 2, tailcone on.
•-..-.- F l ight
20 - _- - Predicted
lO-
f e p
I I I I J
-i00 2 4 6 8 I0
T i me f r om separat i on, sec
(c) Orbiter pitch attitude9 @o, time history.
.--- Flight
12 - --- P r edicte d
. - _ '_ _,_ _,_ • 8 - _'-_., I I I I
0 2 4 6 8 10
Time from separation , sec
(d) orbiter angle of attack_ _o_ time history.
Figure 13.- Continued.
-- Predicted
No 1
2 ____'_-_- -- Flight
I I I I I
0 2 4 6 8 10
Time fromseparation, sec
(e) Orbiter normal load factor 9 NZo _ time history.
2 - -- Flight
--- Predicted
,N
N ° 1 - - ---"
I I_ I I I
0 2 4 6 8 10
Time from separation , sec
(f) SCA normal load factor 9 NZc 9 time history.
2 - -- Fl i ght
- -- Predic t ed
z % I I I I I 0 2 4 6 8 10 Time from separation , sec (g) Rel a tive norm a l load f a ctor, A H Z_ time history.
Figure 13.- Continued.
50 O Flight (trajectory of
O r biter aft attach point )
P r edicted
j E N
= 3 2 s ee
.2
= .9
•£ 20
2 . 0
I stabi l ize r
1.2
0.8
1 .6 _
. 4 I,
0 I0 20
Relat i velong i tud i nalseparat i ond i stance , _X , m (h) Orbiter aft attach point separation trajectory .
(Origin is fixed at SCA aft attach point,) Figure.13.- Concluded.
3 c . g . = 6 5 . 9 percent
2 8e = 2.5 °
'_ I
AN.z = 0.917g
"_" 0 = 0.956 deg / sec2
o
-_ 0
-i
-2
- .3 I I I I I
0 2 4 6 8 10
Time from separation , sec
(a) Orbiter pitch rate command 9 @ c md_ time history.
4 - ---- Flight
3 , --- Predicted
( J
-_ 0
-1 " "_ %%%_iIi _
-2 - \''_
-3 I I I I I
0 2 4 6 8 I0
Time from separation, sec
(b) Orbiter pitch rate, 8o, time history.
Figure 14.- Free flight 3_ tailcone on.
20 - _ Flight
--- Predi cted
10-
€:b
O-
-I0 I, I I I .i
0 2 4 6 8 10
Time from separation, sec
( c ) Orbiter pit c h attitude_ Oo , time history.
12 - -- --- Flight
- -- Predicted
"o
I I I I I
0 2 4 6 8 10
Time from separation, sec
(d) Orbiter angle of attack 9 C_o , time history.
Figure 14.- Continued.
2 -.--.--. Flight .
---- Predicted N° 1 Z t I I I I I 0 2 4 6 8, 10 Time from separation , sec (e) Orbiter normal load factor_ NZo , time history.
2 - _ Flight
--- Predicted
N Z I I I I I 0 2 4 6 8 10 Time fromseparation, sec (f) SCA normal load factor_ NZc9 time history.
2 - _ Flight --- Predicted N I I I I I I 0 2 4 6 8 10 Time f ro m sepa r at io n , sec (g) Relative normal load factory ANz_ time history.
Figure 14.- Continued.
50 - 0 Flig h t (traj ec t o ry o f
Orbiter aft attac h p oin t)
P red i cted
(h) Orbiter aft attach point separation trajectory.
(Origin is fixed at SCA aft attach point.)
Figure 14.- Concluded.
CA = captive active
12 - FF = free flight
u _; Separationwindow 7
_ n 8 -- 12 0
g 4 - CA3 FF1
of5 °
Target FF2
o FF3
",, i t1
0 - r_
_ -4-
o
-8 I I I I I I
0 .2 .4 .6 .8 i .0 I .2
Relative normal load factor , A N z , g
Figure !5.- Orbiter tailcone-on separation initial conditions.
(See figure 2.)
_i _ __ Figure 16.- Free flight 3_ SCA / Orbiter tailcone-on separation trajectory (2 frames per second).
12- 8 -- "O
:2
4- o A ",._
©
Target o o 0 - x : B
-
o -8 '" I I I I I I 0 .2 .4 .6 .8 1.0 1.2 Relative normal load factor, ANZ , g (a) Relative normal lo a d factor compared to Orbiter pitch acceler a tion.
z .0.6 t- O S eparation windo e-- . 2 A • t.- I I I I --_ 0 . 2 . 4 . 6 . 8 1 . 0 1 . 2 0 ,J r,.,, Relative normal 10ad factor, AN z , g (b) Relative normal load fact o r compared to relative longitudinal acceleration.
Figure 17.- Free flight 4 GO / NO-GO separation window. (A and B indicate the load cell data used.)
c.g . = 66.25 percent
= 7 °
e _l_ z - J..U__ / g
= -0.049 deg / sec
3 o
o 2
I
GJ
-1
- 2 I I I I I
0 2 4 6 8 i0
Time from separation , sec
(a) Orbiter pitch rate command_ Ocmd_ time history.
3 - _ _ _ Flight
2 ,, ,_ . "-"" _-'..-_a ' " ' _,. --- Predicted
. o
"Cl:_
-1
-2 n I i i i
0 2 4 6 8 i0
Time from separation, sec
(b) Orbiter pitch rate9 0o_ time history.
Figure 18.- Free flight 4_ tailcone off.
5 0 - ---- Fl i ght
.....P r ed i cted
-10 I I t I I
0 2 4 6 8 10
T i me from separation, sec (c) Orbiter pitch attitude_ @o, time history.
-- Flig h t o
4 -
. I I I I I
0 2 4 6 8 I0
. Timefrom separation, sec
(d) Orbiter angle of atta ck 9 (_9 time his t ory.
Figure 18.- Con t inued.
o 1
N Z I I I I I
0 2 4 6 8 10
Time fromseparation , sec
(e) Orbiter normal load factor , NZ o _ time history.
2 - -- Flight
--- Predicted
N Z _ S_
I i I I I
0 2 4 6 8 10
Time fromseparation, sec
(f) SCA normal load factor_ NZc 9 time history.
2 - -- Flight
- -- Pre d icted
N 1
Z I I I I l
0 2 4 6 8 10
Time fr o m sepa r ati o n , sec
(g) Relative normal load factor , _N Z , time history.
Figure 18.- Continued.
40 0 Flight (trajectory of
Orbite r aft attach point)
-- Predicted
E
N 30
<J
o o ,m
•_ 2- 0
Se e >
> SCA vertical stabilizer
- = _I0 1 .6
(D ,.w
" 1 -2
.8
0 10 20
Relative longitudinal separationdistance,AX , m
(h) Orbiter a ft a tt a ch point separation trajectory.
(Origin is fixed at aft attach point.)
Figure 18.- Concluded.
5 - c . g. = 6 6. 25 percent
Be = 7 °
AN Z = l.Og
3 e ' o -i . 0
4 _ deg / sec 2
E O . (| - * :: b -1 -
-2-
-3 I I I I I 0 2 4 6 8 10 Time from separation, sec (a) Orbiter p itch rate c omman d 9 _ c md , time history.
-2 - +
-3 I I I I J
0 2 4 6 8 10
Time from separation, sec
(b) Orbiter pitch rate_ Co, time history.
Figure 19.- Free flight 5_ tail c one off.
20- -- Flight
- -- P r edicted
10-
(_)0 0 *" "" " " '
-10 I I I I I
0 2 4 6 8 10
Time Fromseparation , sec
(c) Orbiter pitch a ttitude , 0o , time history.
12 - -- Flight
--- Predicted
I I I I I
0 2 4 6 8 lO
Time from separa t ion , sec (d) Orbiter angle of attack, _o9 time history.
Figure 19.- Continued.
2 _ Flight .... Predicted "o 1 S I I I I I
0 2 4 6 8 i0
Time from separation , sec (e) Orbiter normal load factory NZo _ time history.
" 2 - -- Flight ---- Predicted I I I I I
0 2 4 6 8 I0
Time from separation, sec (f) SCA normal load factor_ NZc 9 time history.
2 -- Flight - -- - Predicted
I I I I I '
0 2 4 6 8 10
Time from separation, sec (g) Relative normal load factory ANz_ time history.
Figure 19.- Continued.
50 - 0 Flight (trajectory of
Orbiter aft attach point)
Predicted
40 3 " 2 s e c
E
- 0
N
<1
= "9
_ 30
,n
= 3
o ,D
_- 2q
,_
o 20
>
._> 2 . 0
n ,,
SCA vertical stabilizer
I0 1 .6
0.8
1 " 2 0 . 4
0 I0 20
Relat i ve long i tudinal separation distance, AX, m
(h) Orbiter aft attach point separation trajectory.
(Origin is fixed at SCA aft attach point.)
Figure 19.- Concluded.
12 - FF = f r ee fl i ght rv Separa t ion window "-7 o 8-
°
c- o °_ Target o 0 - FF4 O FF5 " °_ o -8 I I I I _ i 0 .2 .4 .6 .8 1 . 0 1.2 Rela L ivenormal accelera L ion , A N Z , g (a) Relative normal acceleration compared to Orbiter pitch acceleration.
z 0.6 - FF = free flight <3 ,_ .4- S epa r at i °n w i nd°w 7j e.- I1 ) _._N,_ .2 - [ get •O F F5FF4 and . _- I I I I I --_ 0 .2 .4 .6 .8 1 . 0 I .2 n. ,.
Re l ative n0rmal accele r ation, ANZ , g (b) Relative normal acceleration compared to relative longitudinal acceleration.
Figure 20.- Orbiter tailcone-off separation initial conditions.
Figure 21.- Free flight 5, SCA / Orbiter tailcone-off separation trajectory (2 frames per second).
A PPENDIX - G RA C IE PRO G R AM Th e G r o und Redu c ed Aer o dynami c C oef f i c ient s and In s trumentation E rror s (GRA C IE) program was developed a s a tool to aid in flight te s t verifi c ation of the Orbiter / S huttle C arrier Air c raft (S C A ) s ep a r a tion aerodyn a mi c data ba s e. The program ca l c ulate s the force and moment c oeffi c ient s of e ac h vehi- c l e in proximity to the other_ using the load me as urem e nt s y s tem (LM S ) dat a _ the flight in s trument a tion dat a (_ B9 body rate s 9 a cce ler a tions9 et c . ) _ a nd " the vehi c le mass properties. The un c ertainty in ea c h c oeffi c ient i s deter- mined 9 b as ed on the quoted in s trumentation acc ur ac ies. ( U nit s of mea s urement are tho s e used in the software de s ign.)
SYM BO LS
[A] tran s f o rmation m a trix to c h a nge f rom SC A b o dy ax i s to Orbiter body axi s c oordinat e s y s tem Io 1 0 I c i o 0 -s in " i1 si n i ° 0 c o si C vehi c l e aerodynami c c oeffi c i e nts C c o e ffi c ient s F vehi c le for c es F load c ell for c e c omponent s [G] tran s form a tion m a trix to c hange from body axi s to s t a bility axi s
Co°°° O o i
Ls n° ° 0oOS oj
[I] vehicle inertia matrix 9 slug-ft 2 I ixxO-Ixz- ] 0 Iyy 0 / -Ixz 0 Izz J io Orbiter incidence angle , deg L attach strut forces as measured by the load meas'urement .
system , Ib vehicle reference length used for calculating vehicle moment coefficients , ft M vehicle moments m _ vehicle mass, s lugs NX, Ny9 NZ linear acceleration a t vehicle center of gr a vity, g p vehicle roll rate, deg / sec vehicle roll acceleration, deg / sec 2 q vehicle pitch rate , deg / sec dynamic pressure, Ib / ft 2 vehicle pitch acceleration, deg / sec 2 R vehicle position vector r vehi c le yaw rate , deg / sec vehicle yaw acceleration, deg / sec 2 S vehicle reference area, ft2 T SCA thrust , ib V velocity, ft / sec W vehicle weight, Ib X , Y , Z rectangular Cartesian coordinates vehicle angle of attack , deg vehicle angle of sideslip , deg y vehicle flightpath angle, deg ANX, _N y , _N Z relative lo a d fac tor s , g e vehicle pitch angle , deg Orbiter instantaneous pitch acceleration, deg / sec 2 tilt angle of forward strut, deg t p center of gravity (c.g.) relative position vector (AX, AY, AZ), ft 0 c.g. to attach strut moment arm, ft vehicle roll angle, deg vehicle yaw angle, deg vehicle angular velocity vector - p , q,r vehicle angular acceleration vector - _,_,_ Subscripts: A axial a aft c SCA vehicle c.g. center of gravity f forward D drag L left N normal o Orbiter vehicle . R right Operator: (') uncertainty in designated coefficient
PROGRAMDESCR I PT IO N AND ASSUMPT I ONS
The GRA C IE progr a m u s e s flight test d a t a to det e rmine a erodynami c c oe f fi - cient s and their c orre s ponding un c ert a intie s for c omp a rison with wind tunnel predi c ted value s . The program manipulates LM S for c e s 9 S C A body motion s 9 vehi- c le configurations , and vehi c le m ass properties to output tabulated and plotted time historie s of Orbiter proximity9 SC A proximity, and m a ted vehi c le a erodynami c force and moment c oeffi c ient s , a s well as relative norm a l lo a d f a c t o r (AN Z ) a nd O rbiter in s tant a neo u s pit c h a cc eler a tion (8o). The LM S dat a , the SC A body motion data, a nd the vehi c le c onfigur a tion a re obtained from a ground-re c orded telemetry d a t a tap e on whi c h a ll instrument a tion ca li- br at ion s ha v e b e en re c orded. The ve hi c l e m ass prop e rtie s a n d the SC A pre- dicted data time historie s are input throug h s ubroutines be ca u s e they require postflight c al c ulation s a nd are not re c orded on the dat a t a pe.
T he program perf o rm s thre e b as i c oper a tion s us ing the flig h t test d at a .
The equations of motion and the aerodynami c un c ertainty c al c ulation s are m a de with data retrieved from flight te s t instrumentation, and the predi c ted values of the c oefficient s are determined. The following s e c tions de sc ribe the s e operation s .
Equ a t ion s of Motion A s a ba s i s for cal cu l a ting equ a tion s of motion, the mated vehicle i s ass umed to be a rigid body in motion with re s pe c t to a fixed c oordinate s y s - tem XY Z (fig. 2 2 ). Affixing a s e c ond s et of axe s to the ca rrier aircraft, with the origin ( c ) lo ca ted at the carrier c .g._ and ob s erving it s motion a llow s ev a luation of the motion of a ny other point in the mated c onfigur a tion 9 namely the Orbiter c.g., a s well as the mated c .g. For example_ the a cc el e r a tion of the Orbiter c .g. (o) c a n be determined by knowing the relative position (p), the li n e ar acc eler a tion ( R ) , angul a r r a tes (_)9 a n d a ngul a r a cc ele ra ti o n s (_) of the ca rrier c.g. (c).
R o = R c+_cx P +c u cx (_cx p )+_ o l c +2_cx Vo l e H owever , the m a te d vehi c le is as sumed to be a rigid body; therefore , the rela- tive velo c itie s and a c celeration s between the c.g. 's are a o / c = Vo / c = 0 Therefore9
_o=_ c . , ; , o× p . ,, , c ×(,, ,c + p)
The t otal result a nt or a p p lied f o r c e s o n ei t her vehi c le are t hen
F =m c R c
c tota l app l ied oo Fo = m ° R o total applied and F = tomB rntotal m applied Simil a r use o f kinematics provides the equations for cal c ulating the re s ultan t moments ( M ) on each vehicle 9 i.e.
M = I t ] c& _ +% x [ I ]_
Ctotal applied IVl = [I]o_o+_o x [I]o_ tota_ o applied and M = [I] m O J-+ O Jm X [I] m_m mtota I m applied where _o = [A] e Jc' _o = [Alc_ an d From figure 2 3, th e lo a d cell o u tput s , expr e ssed in th e c a r r ier b ody axis co - ordinate s y s tem, are as .follow s: , Ffy = forward side force Ffz = f o rw a rd vertic a l for c e ( p ar a llel to s trut a xis) Ffx7_ Ffz 7 drag and vertical component s of forw a rd verti c al strut force (carrier body a xis coordinate system) F L x = left a f t d rag f or c e FLz = left aft verti ca l for c e FRx, = right aft drag for c e " FRy = right aft s ide force FRz = right aft vertic a l for c e where Ffx 7 = Ffz sin ;k, Ffz7 = Ffz co s and 1 723 336. 5-- 1 72 3333. 7 c o s(i o + 2 . 7 34 ° X= 88 . 27o sin-1 f 9 29'098sin(i° +2"734°) ). 1 Al s o s hown in f igure 2 3 are th e mom e nt ar m s ( 1 1 ' 1 2'" . ., l 7 ) f rom th e Orbiter c .g. to ea c h lo a d cell att a ch point9 based on th e c arrier body axi s coordi- nate s ystem. Using figure 23 in c onjun c tion with figure 24, the mom e nt arm s a re determined from the following relations_ noting that the att ac h point location s are in the Orbiter body c oordinate system :
_ z_ ' g ' ° -z _______ f _=tan - ' X°'°'o -- X' J
Ca= tan XR --Xc'g ' o / If _ z c ' g'°- c.g 'o # _'a = sin q _a tZ --ZRt l , = I a sin( q 5 a + io) / 12 1 2 = I a c°s(_ a . io_ / 1 2 _'3: 1 ' sinef -io) / 12 1 4 = / f cos(_ b f -- io) / 12
_0-- -(_,+ _0 0.o) / ,_
_0: (_,-_c.0.o) / ,, F rom figure 2 5 , and using the Orbiter mom e nt arms pr e vi o usly ca l c ulated 9 the position ve c tor is N oti c e that in f igure 2 5, the a tta ch point lo c ation s a re in the carrier body a x is c oordinate s ystem.
The following free body di a gram s a nd c orre s ponding equation s of motion are u s ed in c al c ulating the aerodynami c for c e and moment c oeffi c ient s of th e mate d v e hi c le_ the SCA in proximity to th e Orbit e r_ and the Orbit e r in proxim- ity to the SC A.
Mated vehicle aerod ynamic coefficients .-For c e coeffi c ient s (drag_ s ide for c e_ lift) Ftotal Fthrust _ I R c/I Rm ,11 Fto t a I -- F aero + F thrus t applied applied Faero = mm Rrn -- Fthrus t F aero C body _lSc axis C stability = [G]cC bo.dy ax i s ax i s "[G] c = Transformation matrix to changefrom carrier body axis to carrier stability axis Mate d vehicle aerodynamic coefficients . - Mom e nt c o eff i c ients (rolli n g moment_ pit c hing moment 9 yawing moment) M total = M aero + M th r ust applied M total = [I]m_c+(_c x [I]m_ c applied M ae r o= M tota I -- M thrust applied M a e r o C m o men t - _Scl c Carrier aerod ynamic coeffic ients ( p roximity). - For ce c oef f i c ient s (dr a g9 s ide f or c e 9 lift) F l o ad cel lf Fl o a d FtotaI Faer° cella Fthrust R Ftotal =Faero +FIoad +Fthrust ap p lied cell F tota l = m c rtc applied F l oa d = Lc = L f + LL+ L R ce ll F aer o = m cRc -- F l oa d -- F th rust ° cell Faero C b o.d y = _ S c a x is C stabilit y = [ G] c Cb o.d y axis axis Carrier aerody namic coefficients (proximity) . - Mom e nt c o eff i c i e nt s (r o lling m o m e nt 9 pitching moment 9 yawing moment) M total =Macro +MIoad + M thrust applied cell M to t a I = [ I ] c _ c + c o c x [ I ] cO _ c applied c ell s =l c Mac ro= Mto t a I -- M l oa d --Mthrust a ppl ied ce ll M aero C moment - S c / c 7O Orbiter ae rod Tnamic co efficients (proximity) .- F or ce c o ef fi c ien t s (d ra g _ side force_ lift) Ft o t al F a e r o appli e d Fl°a d !
cellf Fl oad cell a R R o ¢ Ftotal =Fa e ro + F I oad a pp l ie d cell F total :m oR o:m o [ Rc+ _ c × p +OJc x ( OJc × p )_ applied Flo a d= L o = L f+ L L + L R c ell F a er o = [A](mo R o -- Lo ) Faero C body - a xis q So Cstabil i ty = [G] o C body a x is ax is [A] = Tr a n s form ati on from carr i er t o Orb i ter c o ord i nate sy s t e m at inc idenc ea ngle i° O rbiter a erodyn amic coefficients (proximity) .- M oment coe ff i c i e nts (rolling moment _ p it c hing moment 9 yawing moment) M total = M aer o -- M load ap p lied cell M t o ta I = [I]o _ o +O J ° X [I]oO J ° appl i ed
Mo0 E ( . o Lo )
cell s = l M aero = M total -- Mio a d applied cell M ae ro C m o m ent - S olo Orbiter pitch acceleration: M _ aeroy I YYo Re l a t ive l oa d fac t o r s : / Wo + W c _ /_ N z = Lo _ - "_f W' _ " _ ) A e rodyn a mi c U n c ert a inti es An integr a l p a rt o f the separ a tion a nalysi s i s knowing the un c ertainty a ss ociated with ea c h c oeffi c ient a nd how that un ce rtainty affe c t s the size of the sep a r a tion window as well as the vehi c le traj ec tory. E a c h a e rodynamic coefficient is a function of i independent measurements_ ni9 and the uncer- tainty of ea c h me a surement is _n i.
C = f(n 1 , n2, n3..... ni) (1) The uncertainty in each calculated coefficient is obtained by using the fol- lowing equation."
[coc_ _+ [ ,c_ _ +/ , c_ _ _] ' ' _
_c =L\_j (An1) \_n2 ) (An2) +. .. \ 6 ni / (Ani) (2)
T he un c ertaintie s in the aerod y n a mi c c oeffi c ient s are b ase d on the quoted a c cura c ie s of the load measurement s ystem a nd the flight te s t in s trument a - tion.
The un c ert a inty in the Orbiter for c e and moment c oeffi c ient s are ca lculated as follows. T he Orbiter a erodynami c for c e s a re f ir s t c al c ul a ted with re s pe c t to the SC A body c oordinate sy s tem from the following e quation s:
_o(Nx +_Z-_Y+..._Y -q_X + , _Z- ,_X) -_.F x
Cx = _S° _Fy mo(_Ny__ --ISAZ+_AX--p2Ay+pqAX+rqAZ-r2Ay ') _- Cy = _So mo(N z +I_AY _AX-- p2AZ + prAX-, q2AZ + qrAY) - Y ,F z C Z = _So From equation (2)9 the uncertainty in CX is mo(qA y + rAZ)t2 mo(PAY-- 2qAX)- 1 2
q '= L 0So J (Aq)2
15 ' = n eg l ig ibl e mo. 'k Z_2 ( _moZ_Y_2
_= k_-Wo ) (_)_
mo(N x + _l'..%Z - _JY - qp_Y - q2AX + rp&Z - r2AX) + L x
_'= 0_)_
q 2 S o
, : ( 'Y/_F, _2
F'x \_s U k "x }
ACx=( N_( + p' + q ' _ r ' + 6' + _1 ' + _ ' + _' + F'fx + 'FLx +F'Rx ) _I / 2 The uncertainty in CZ is calculated similarly. The coefficients are then transformed into the Orbiter body axis coordinate system.
C A = C x cos i° - C z sin i° o CN = Cx sin i ° + Cz cos i° o The uncertai n ties in these two coefficients are ACAo = I(COSio)2 (ACx)2 + (sin io)2 (ACz)211 / 2 ACNo = [ (sin io)2 (ACx)2 + (cos io)2 (ACz')2] 1 / 2 Fin a lly 9 these coeffi c ients are transformed into the Orbiter stability axis coordinate system CDo = CAo COS (a ' c + i°) + CNo sin(_c + i° ) Clift ° - CAo sin((_c+io) +CNoCOS(_+io) and the uncertainties in the Orbiter coefficients of lift and drag are
_c,, , . o _-t[o0._ o o +,o) ] '(_C_oy +[., o ( o o +,o) ] =( . C.o ) _
_2 2 _1 / 2 + [CA° sin(c t' c + i°)- CN° c ° s(C { c +i°) J ( /k _' c ) I
.Coo: t[.,° ( oo + , o)] _(.C_o)_ + [co. ( oo +,o ) 3_ (_C . o)_
+ [ CAoC°S( O _c+io) +CNoSin(%+io) ] 2(A(_c)2fl / 2 T he un c ert a inty in t h e Orbiter s i d e for c e coe f fi c i e nt_ A C Yo_ i s found in the sa me w ay.
The Orbiter moment coeffi c ients are based o n the following equations :
_o,xx + % . o(,ZZo-,_o) + ,X.o q O+.O%)
C = m x qSobo , , -- Ffz pfy + Ffypfz - FLZPLy -- FRzPfy + FRyPRz + " . , qSob o
_o,_o + _o ,o (,XXo- ,Z_o)+,XZo (,o_- %_)
Cm y qSoC o Ffz p f x -- Ffx P fz + FLZ P LX -- FLX P LZ + FRZ P RX -- FRX P RZ + _So_ - o _ol zzo + Poqo (IYYo - Ix×_- I×zo (qoro - I_o ) Cm Z = qSo bo - Ffy p f X+ Ffx P fy +FLx P LY-- FRy p Rx +FRx P RY + qSobo Again u s ing equation (2), the uncertainty in the Orbiter pitching moment i s I 2 "
\_So---g ! (_)_
I XX ° -- I ZZ ) + 2Po l XZol 2
p ' = _So_ J (_p) 2
.o(,xxo ,_Zo)_ .o,XZo (_ , )_
E t2
r ' = qSoc XZ o _l = .
_2S o _" ClolYYo + PoroQXXo-IXZo) +1 (ro2--po2) Ffz Pfx -- Ffx Pfz +FLz P Lx -- FLx PLz + FR ZPRX -- FRX P RZl 2 + _2S o E - j ( A_i)2
{ P fx ,_2
E ; z : _To_ ) (AFfz) 2 _. P fz . % 2
_x : _-_o_ ) ( _ ' x)_
{ P L x. % 2 F'LZ =[_) (AFLz) 2
__ ( _ L _ _
_'_x \_so_ / (_'x) =
_. P RX_ 2 FRZ= [_ ) (AFRz) 2
F_x \_So -- - 7 /
t A C my + p , + r, + _, + F_Z + F , fx + F,LZ + F,LX + FRZ + FRX 1 / 2
=(_ )
The uncertainties in the r o lling m oment and the yawing m om en t are calculated similarly.
Analysis of the uncertainties in the SCA proximity and mated vehicle co- efficients is performed in a like manner.
X c . g . o 0 rb i te r I I c . g For wa rd z atta ch Lf c ' Q '° poi nt 1 3 Sf , w Zf_ L a I 1 4 Aft , xf i i I attac h C a point Z L ,Z R XL , X R Figure 24.- Orbiter attach point moment arms.
0 rbi ter cg Aft attac h ' £1 point I £2 , ,=
Carr ie r
/ c.g. I
I
Z _ _ I c ' g ' 7 4 7 I I . X XR c-g.747 Figure 25.- Relative c.g. locations.
1 . Report No. 2 , Government A c c essio n No. 3. R e cipient ' s CatalogN o , NASA TM-58223 4 . Tit i e and Subtitl e 5 . Report Date ORBITER / SHUTTLE CARRIER AIRCRAFT SEPARATION: WIND TUNNEL, May 1980 SIMULATION, AND FLIGHT TEST OVERVIEW AND RESULTS 6. P erfor m ing Organization Code JSC-16711 7. Author(s) 8, Performing Or ga nizationR e portNo .
D. J. H o man, JSC, and D. E. Denis o n and K. C. Elchert, Rockwell International 1 0. Work Unit No .
9 . Perfor m ing Organization NameandAddress 9 5 3-36-00-00-72 _= Lyndon B. Johnson Space Center 1 1.C o ntract or G rantNo .
Houston, Texas 77058 1 3. Ty p e of Repo rt and Period Co vered 1 2 . Sponsoring Ag e ncy Nameand Address Technical Memorandum Nati o nal Aer o nautics and Spa c e Administration i 4. Sponsoring AgencyCode Washington, D.C. 20546 15. Supplem e ntary Not es 16. Abstra c t A summary of the a ppr o ach and landing test phase o f the Space Shuttle Pr o gram is given from the Orbiter / Shuttle Carrier Aircraft separation point of view. Material in the report covers the data and analyses used during the wind tunnel testing, simulation, and flight test phases in preparation for the Orbiter approach and landing tests.
Predicted separation parameters are compared to actual flight data.
1 7. Key Words(Suggest ed by Author(s)) 1 8 . DistributionStatement Test flights Free flight , J Space Shuttle Active-captive flight STAR Subject Category: Orbiter Captive-active flight 18 (Spacecraft Design, Testing , i Wind tunnel Approach and landing tests and Performance) Simulations l g . S e cu r ity _ a m i f . (o f this r e _ rt) I 2 0 . S e cu r itv C lass if .(o f t h is _) ] 2 1 . No , o f P a_+ + 2 2 . _ic++ Unclassified Jl Unclassified I 87 1 $5.00 +F o r _ le by th e N a tio n a l Tech n i _ l In f or m atio 6_ r v ice 0 : Sp r i ng fie l d , Vir g in i a 2216 1 JS C Fo rm 1 4 24 ' R _ Nov 7 5 ) NASA -- - _C U. $ . G OVERNME N T P R I NTI N GOFFI C E:198 0-- 6 7 1+09 _ 1610 f_ .i'