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General Disclaimer One or more of the Following Statements may affect this Document This document has been reproduced from the best copy furnished by the organizational source. It is being released in the interest of making available as much information as possible. This document may contain data, which exceeds the sheet parameters. It was furnished in this condition by the organizational source and is the best copy available. This document may contain tone-on-tone or color graphs, charts and/or pictures, which have been reproduced in black and white. This document is paginated as submitted by the original source. Portions of this document are not fully legible due to the historical nature of some of the material. However, it is the best reproduction available from the original submission. Produced by the NASA Center for Aerospace Information (CASI)

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NASA TM X-56039 STATUS OF THE NASA YF-12 PROPULSION RESEARCH PROGRAM James A. Albers N76-1Q152 (NASA-Tr.-X-56039) STATUS OF 11fiT N4SA YF-12 (NOSA) 32 p HC PROPULSTON ?ESEkRCH PLOGR41 $11.00 March 1976 CSCL 21F Unclas G3/0 7 20679 NASA high -number Technical Memorandums are issued to provide rapid transmittal of technical information from the researcher to the user. As such, they are not subject to the usual NASA review process. C a Nip 6a^`NLHI Dryden Flight Research Center Edwards, California 93523

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I Report No. TM X - 56039 4 Title and Subtitle 2. Government Accession No 3. Recipient's Catalog No, 5. Report Date March 19 ,6 STATUS OF THE NASA YF-12 PROPULSION RESEARCH PROGRAM 7. Authorls) James A. Albers 9. Performing Organization Name and Address Dryden Flight Research Center P.O. Box 273 Edwards, California 93523 12 Sponsoring Agency Name and Address 6. Performing Organization Code 8. Performing Organization Report No H - 935 10. Work Unit No. 516-51-02 11. Contract or Grant No 13 Type of Report and Period Covered Technical Memorandum National Aeronautics and Space Administration Washington, D.C. 20546 15 Supplementary Notes 16 Abstract 14 Sponsoring Agency Code The YF-12 research program was initiated to establish a technology base for the design of an efficient propulsion system for supersonic cruise aircraft. The major technology areas under investigation in this program are inlet design analysis, propulsion system steady-state performance, propulsion system dynamic performance, inlet and engine control systems, and airframe/propulsion system interactions. This report discusses the objectives, technical approach, and status of the YF-12 propulsion program. It discusses the results obtained to date by the NASA Ames, Lewis, and Dryden research centers. The expected technical results and proposed future programs are also given. 17 Key Words (Suggested by Author(s) l Supersonic aircraft Aircraft propulsion Inlets Inlet engine compatibility Aircraft controls 19 Security Classil (of this report) Unclassified 7T^ 18. Distribution Statement Unclassified Unlimited rity Classif (of this page) 21 No of Pages 22 Price' assified 31 "For sale by the National Technical Information Service, Sprin (Ifield, Virginia 22151

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CONTENTS INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 OBJECTIVES . . . . TECHNICAL APPROACH . . . . . . . . . . . . . . . . . . . . . . . . . . 3 PROGRAM STATUS . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Inlet Design Analysis . . . . . . . . . . . . . . . . . . . . . . . . . 3 Propulsion System Steady-Stato Periormance . . . . . . . . . . . . . . 3 1/3-scale test results . . . . . . . . . . . . . . . . . . . . . . . 3 Full-scale test results . . . . . . . . . . . . . . . . . . . . . . . 4 Flight test results . . . . . . . . . . . . . . . . . . . . . . . . 4 Wind tunnel/flight comparisons . . . . . . . . . . . . . . . . . . . 4 Propulsion System Dynamic Performance and Inlet-Engine Compatibility . . 5 1/3-scale test results . . . . . . . . . . . . . . . . . . . . . . . 5 Full-scale test results . . . . . . . . . . . . . . . . . . . . . . . 5 Flight test results . . . . . . . . . . . . . . . . . . . . . . . . . 5 Wind tunnel/flight comparisons . . . . . . . . . . . . . . . . . . . 5 Inlet and Engine Control Systems . . . . . . . . . . . . . . . . . . . . 6 Inlet control system . . . . . . . . . . . . . . . . . . . . . . . . 6 Turbine inlet gas temperature control system . . . . . . . . . . . . 6 Stability bleed system . . . . . . . . . . . . . . . . . . . . . . 6 Cooperative control system . . . . . . . . . . . . . . . . . . . . . 6 Airframe/ Propulsion System Interactions . . . . . . . . . . . . . . . . 7 1/12-scale test results . . . . . . . . . . . . . . . . . . . . . . . 7 Flight tuft studies . . . . . . . . . . . . . . . . . . . . . . . . . 7 PROGRAM PLANS AND SCHEDULES . . . . . . . . . . . . . . . . . . . . . 7 Present Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Compressor face dynamics . . . . . . . . . . . . . . . . . . . . . 7 Boundary layer dynamics . . . . . . . . . . . . . . . . . . . . . 8 Atmospherically induced turbulence . . . . . . . . . . . . . . . . 8 Inlet transients . . . . . . . . . . . . . . . . . . . . . . . . . . 8 rative control system . . . . . . . . . . . . . . . . . . . . . 8 ed Technical Results . . . . . . . . . . . . . . . . . . . . . . . 8 ion system steady-state performance . . . . . . . . . . . . . 8 iii

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ion system dynamic performance . . . . . . . . . . . . . . . 9 Control systems . . . . . . . . . . . . r . . . . . . . . . 9 Airframe /propulsion system interactions . . . . . . . . . . . . . . 9 Proposed Future Programs . . . . . . . . . . . . . . . . . . . . . . . 9 . . . . . . . . . . . . . . . . . . . . . . . . 9 Propulsion test bed Advanced shock sensing techniques . . . . . . . . . . . . . . . . . 9 Shock stability bleed system . . . . . . . . . . . . . . . . . . . . 10 Flight effects of annular nozzle noise suppressors . . . . . . . . . . 10 Airplane performance and drag prediction . . . . . . . . . . . . . . 10 CENTER RESEARCH AREAS . . . . . . . . . . . . . . . . . . . . . . . . . 10 . . . . . . . . . . . . . . . . . . . . . . . . . . 11 REFERENCES . . . . . iv

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STATUS OF THE NASA YF-12 PROPULSION RESEARCH PROGRAM James A. Albers Dryden Flight Research Center INTRODUCTION Supersonic transport aircraft require propulsion systems that operate efficiently through a wide range of altitudes, and at speeds from subsonic to high supersonic cruise. To avoid penalties in engine size, weight, snd fuel consumption, the inlet system must be designed to supply air at the maximum pressure with minimal drag and interference. The inlet system must also be able to match the airflow requirements of the engine over a wide range of flight conditions. To optimize an inlet fur a given aircraft mission requires extensive tradeoffs between performance at design and at off-design conditions. A first step in the o ptimization of the propulsion system is an analytical study of the various inlet geometries that match the engine requirements. This is followed by wind tunnel testing of scaled models prior to flight testing. In general, conditions in the wind tunnel do not exactly duplicate flight conditions. With scaled models. the Reynolds numbers and the local flow field (to not always correspond to those in flight. In addition, the geometry and the instrumentation location and accuracy of wind tunnel models are difficult to match to those of the flight hardware. Since the flight hardware and its expected performance are determined from scaled wind tunnel models, scaling techniques that allow the extrapolation of subscale inlet data to full-scale flight are necessary. Many of the current propulsion system problems for supersonic cruise aircraft involve inlet-engine compatibility. Insufficient propulsion system stability margin caused by pressure distortion has been and continues to be a significant problem in aircraft development. It is presently not clear l.ow dynamic data from model tests can be used to predict the stability margin of the propulsion system in flight. Another area of major concern to the propulsion system designer is the prevention of the inlet unstarts , which result when the terminal shock moves out in front of the cowl lip. Unstarts can occur when either internal disturbance or external

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disturbance occurs in flight. New propulsion control concepts are needed to position the terminal shock in the inlet duct. At present, mixed-compression inlets have variable geometry features that are programed by a variety of engine, inlet, and airframe variables. For example, in the YF-12 inlet, variable bypass doors and a spike or ramp move as functions of Mach number, angle of attack, normal acceleration, and angle of sideslip. New stability bleed systems and shock position sensors are required to improve the response of the present control system. Experience to date with supersonic cruise aircraft has indicated strong inter- Actions between the propulsion system and the flight control sy stein . These effects have been traced to the porting of bleed and bypass flows overboard around the nacelle. This can result in separated flow on the external nacelle and in the base and boattail region surrounding the engine exhaust. Thus, the nacelle flow interactions of supersonic cruise aircraft require further investigation. An integrated overall aircraft control system is needed to minimize the undesirable interactions of the inlet, engine, and airframe contnil systems. To establish a technology base for the design of fin efficient propulsion system for supersonic cruise aircraft, a propulsion research program using the YF-12 airplane was initiated in 1969. This report discusses the objectives , technical approach, and status of this propulsion program. It discusses the results obtained to date by Ames, Lewis, and Dryden research centers. The expected technical results and proposed future programs are also discussed. OBJECTIVES The primary objective of the YF -12 propulsion program is to establish a technology base for an efficient inle' system (fig. 1) for supersonic cruise aircraft. The major technology areas under investigation in this program are inlet design analysis and prediction techniques, propulsion system steady-state performance, propulsion system dynamic performance and inlet-engine compatibility, inlet and engine control systems, and airframe/propulsion system interactions. The objectives of the program are as follows: to develop analysis techniques in order to optimize inlet geometries and bleed systems for mixed--compression inlets; to evaluate the effects of Reynolds number, E-aling, flow field, an-1 other wind tunnel and flight differences on propulsion system performance; to evaluate the overall steady-state inlet performance and to determine the operation range of the inlet for various geometries aril flow conditions; to develop scaling techniques that permit the extrapolation of subscale inlet dynamics to full-scale flight; to evaluate the effects of high frequency now fluctuations, or transients. on the stability of the propulsion system; to evaluate new control concepts and stabilization techniques for a mixed-compression propulsion system; to measure and evaluate the effects of atmospherically induced turbulence on the dynamics of mixed-compression inlets; to develop dynamic pressure sensors and other instrumentation for propulsion system testing; to investigate the causes of airframe/ propulsion system interactions; and to utilize the YF-12 airplane as a test bed to investigate new propulsion system concepts, such as turbofan ramjet and variable cycle engine concepts.

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TECHNICAL APPROACH With the use of specially developed high temperature instrumentation, steadystate and dynamic inlet performance was measured. Comparisons are being made between flight data and data from 1/3-scale and full-scale wind tunnel test of the inlet. Methods of extrapolation from wind tunnel ;o full scale are being developed. A comparison of inlet configurations and test facilities is shown in figure 2. A schematic of the YF-12 duct pressure instrumentation is given in figure 3. Wind tunnel and flight evaluations of advance propulsion control concepts, including inlet shock stabilization, ire being made. An example of new control concepts presently being studied is Lewis' stability bleed system, which is shown in figure 4. An integrated propulsion/flight control system (YF-12 cooperative control system) is being developed and is to be evaluated for its usefulness in flight. A conceptual schematic of this integrated control system is shown in figure 5. The interactions between the airplane and the propulsion system could be studied by comparing the flight data with 1/12-scale airplane model data. Some results of a tuft study which show the local flow around a nacelle in flight are shown in figure 6. These experimental results are to be used to refine the analytical models and propulsion simulations. The YF-12 aircraft is to be used as a test bed for the evaluation of future propulsion concepts. PROGRAM STATUS Inlet Design Analysis At the time that the YF-12 propulsion program was initiated, limited design analysis was done in direct support of the program. However, through research programs at Ames and Lewis, a technology base for supersonic inlet design computations has been developed. The inviscid flow analysis of supersonic inlet now fields (refs. 1 to 3) uti l izes the method of characteristics. A boundary layer study of an inlet that uses it bleed system designed for Mach numbers of 2.5 and below is given in reference 4. A bleed study on flat plates is given in reference 5. A viscous flow analysis that is presently being developed to design inlet diffusers is discussed in reference 6. Design analysis for mixed-compression inlets for Mach numbers greater than 2.5 has been done at Ames and sponsored by Ames. Some recent con tract work that has direct application to design bleed systems for supersonic inlets for Mach numbers greater than 2.5 is reported in reference 7. An evaluation of this analytical technique is given in reference 8. 'These analysis techniques could be used to design new inlets for supersonic cruise aircraft. A promising high Mach number inlet has been tested and reported in reference 9. `Recent analysis indicates that this inlet can be further developed by using the analytical bleed design methods used in reference- 7 and 8. In addition, the capability now exists to evaluate inlet now fields at moderate angles of attack (ref. 10) . Propulsion System Steady-State Performance 1/3-scale test results. - A 1/3-scale model of the YF-12 inlet (fig. 7) was tested at Ames in the Unitary Plan Facilities at Mach numbers from 0.9 to greater than 3.

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and at Reynolds numbers (bated on inlet diameter) between 5 X 10 6 and 7 X 10 6 . The aircraft inlet internal geometry was completely simulated from the centerbody tip to the engine face station, including the variable forward and aft bypass doors and the centerbody and cowl bleed systems. The tests involved the investigation and correlation of the steady-state inlet parameters b(Aween th — rirplane snd the model and the investigation and assessment of the effects of scale on the frequency and amplitude phenomena of inlet dynamic turbulence. Hence, a large amount of steady-state and dynamic instrumentation was incorporated in the model. The basic data (including mass flow rates, pressure recovery, distortion parameters, boundary layer surveys, and dynamic data) have been obtained, and the steady-state results are completely documented in references 11 to 13. The dynamic information has not been reduced or analyzed except for a limited statistical analysis reported in reference 12; all dynamic information is recorded on magnetic tapes that are stored at Ames. Analysis of these data and correlations with flight data are discussed in a subsequent section of this document. Full-scale test results. - An extensive steady-state investigation has been completc.-t in T,r wis' 10-Foot X 10-Foot Supersonic Wind Tunnel (fig. 7) . The wind tonne: histallation is described in detail in reference 14. Inlet performance maps were obtained for arious attitudes and angles of attack and several aft bypass door settings . Compre^ 3o^, face total pressure profiles. boundary layer profiles, static pressure distribut; ins , and bypass calibrations were also obtained. The results are given in reference:. Ii and 16. As part of an effor: to obtain accurate airflow measurements in flight, an engine airflow calibration was 1 erformed tit Lewis Research Center's Propulsion Systems Laboratory. The engine L at was installed in the aircraft for the research flights was calibrated with distortion screens that produced distortion patterns that simulated flight conditions. Over 3 percent degradation of performance in corrected engine airflow was observed from the previously used airflow characteristic curve that represents an average engine with little or no distortion. The results of these tests are discussed in reference 1, . Pi ht test results. - The flow sensing probe on the nose boom of the YF-12 airplane and the flow sensing tip of the YF- 12 inlet centerbody were calibrated in the wind tunnel to insure flow conditions comparable to flight (ref. 18) . Flight results for the local flow at the inlet spike tip over a wide range of flight conditions are given in reference 19. Local flow angularity, Mach number. impact pressure, and mass flow at the inlet spike tip are compared with free-stream values. Detailed descriptions of the YF-12C airplane, propulsion system, and instrumentation for propulsion research flights are given in references 20 to 22. Flight tests of the propulsion system included an investigation of off-schedule inlet operation. Some preliminary results are presented in reference 23. Analysis of the steady-state performance of the propulsion system is in progress. Additional flight test results can be found in reference 24. Wind tunnel/flight com arisens - To obtain a meaningful comparison between wind tunnel and flight data, similarity must exist in inlet geometry, test conditions, and instrumentation. References 25 and 26 discuss in detail the instrumentation requirements for flight -to-wind-tunnel comparisons. Preliminary results of comparisons between wind tunnel and flight data are given in reference 27. More recent

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comparisons of wind tunnel tests of 1/3- and full-scale models are made in reference 28. The performance comparison indicates that wind tunnel results can be a satisfactory indication of performance in flight if inlet conditions are matched. Further steady-state data are now being analyzed. The YF-12 overall propulsion system inlet performance is discussed in references 29 and 30. Propulsion System Dynamic Performance and Inlet-Engine Compatibility 1/3-scale test results. - Large amounts of dynamic pressure data were acquired from the tests fit Ames. These data included a 40-tube total pressure survey at the engine face. duct wall static pressures, and boundary layer total pressure rake measurements. The applicable data were stored on magnetic tape. Dynamic pressure data and a statistical analysis of selected pressure data are given in reference 12. Full-scale test results. - The dynamic pressure data included a 24-tube total pressure survey at the compressor face which was identical to the installation for the flight tests. These data were recorded on magnetic tape for comparison with dynamic data to be obtained in future flight tests. Flight test results. - Dynamic flight tests were performed that included deliberately induced unstarts and compressor stalls. An investigation of pressure data during an unstart is being performed to see if an unstart-induced stall is present in flight. Examination of dynamic data indicates that inlet-engine compatibility is good over most of the flight envelope (ref. 23). Since external disturbances can affect the stability margin of the propulsion system , the atmospheric effects on the inlet sy.:tem were investigated. Some preliminary dynamic flight distortion data were reduced on contract. The obje..::ve of the fi • st phase of the work was to integrate the necessary hardware and computer software to make dij,ital data processing possible. The objective of the second phase was to screen inlet dynamic data to extract those time slices during which engine face pressure distortion was most critical and to provide digital records of those events. In addition, duct pressure data were obtained daring transient conditions. The data reduction method is operational, and sonic digital tapes of a limited amount of compressor face and duct pressure data are now being analyzed with a statistical program. Wind tunnel/flight comparisons. - The currently availahle flight dynamic data cannot be compared with wind tunnel data because of zero shifts in the pressure data due to the severe temperature and vibration flight environment. The present wiring installation uses flexible wire that apparently allows zero shifts when the aircraft is operating in this severe environment. This wiring is being replaced with a steel-encased two-conductor cable which should eliminate the problem. After this installation, it s;iould be possible to compare flight dynamic data directly with wind tunnel data.

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R Inlet and Engine Con t rol Systems Inlet control system. - The full-scale YF-12 flight hardware with the duct pressure rat, ,) inlet control system wits tested in the Lewis w ; nd tunnel (ref. 31). These data served as a baseline for coi.tparison with other shock position control systems. This investigation also demonstrated that it digital computer could be used to control a flight-type inlet and could provide P11 the schedules and other complexities required of an actual aircreft inlet control. Frequency response and transient testing of various experimental shock position controls are given in reference 32. Optimum shock position controllers of the proportional plus integral form were determined analytically and tested experimentally. Open loop dynamic wind tunnel data are given in reference 33, which evaluates the response of the flight inlet to internal airflow perturbation. External disturbances are difficult to simulate in the wind tunnel and are better investigated in flight. An electronic terminal shock positiori sensor was tested, and the results tire given in reference 34. Turbine . Met ads temperature control system. - Measurement of the turbine inlet temperat u.eis useful in engine control because this temperature limits the maximum pertormance of the engine. This temperature has been unobtainable, because thermocouple and thermocouple support materials can not withstand a high temperature environment for long periods of time with the high reliability required for use in a control system. Hence, NASA, in conjuration with the Air Force, contracted the development of a high-response fluidic sensor for n turbine inlet gas temperature (TIGT) control system (refs. 35 and 36) . The flight evaluation of this control system is complete and the data are presently being analyzee.. Stability bleed system. - A control concept that is presently being studied at Lewis is a stability bleed system that has been applied to the YF-12 inlet (fig. 4). Two sets of mechanical relief valves are arranged in the cowl. The hleud bypass opening is controlled by poppet valves and is scheduled as a function of the differential pressure between the bleed chamber and the reference volume, An orifice restricts the flow from the back side of each valve and thus dampens the effects of pressure fluctuations (ref. 37) . The optimization of the cowl bleed hole pattern and steady-state testing of this stability bleed system has been completed. Dynamic testing of the valves in the YF- 12 full-scale inlet at Lewis is also complete. Preliminary results indicate good performance of the valves, with increased stability mar- 2.5 to 2.8 range (refs. 38 and 39) . lin for operation in the Mach Cooperative control system. - The cooperative control system as described in reference 40, is mechanized around a digital computer (fig. 5) tnat contains the control laws to process the inputs and generate the outputs. Figure 8 is a flow chart of the various segments of the cooperative control development. A major part of the effort involves the development of the simulation and software. It is planned to incorporate the control integration in logical steps, with optimal integrated control laws not incorporated until late in the program. One airborne digital computer has been purchased, and contro, laws are being synthesized on the simulator. More detailed discussions of this proposed control system are given in references 41 and 42. Other discussions of various aspects of the airframe/propulsion system interactions tire given in references 43 to 47.

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Airframe/ Propulsion System Interactions As part of the YF-12 performance and propulsion program, a limited amount of information was obtained on nacelle flow interaction to lay the groundwork for future flight testing. Surface pressure data were obtained from it modal tested at Ames, and a tuft study of the local flow around the nacelle of a YF-12 airplane was performed. 1/12-scale test results. - Wind tunnel tests of the 1/12-scale model have been run, primarily to obtain force data. A limited number of surface static pressure measurements were obtained during these tests. The pressure orifices were installed on the left wing and nacelle, and measurements were made with various bleed flows through the forward bypass and centerbody bleed louvers. Subsequent testing was done with the same model to obtain loads data. Many additional pressure orifices were installed for these tests. Extensive data were obtained throughout the Mach number range and for various bleed flows. In addition. data for started and unstarted inlet conditions were recorded. Finally, inlet flow field data were obtained by removing the inlet and placing a rotating conical probe in the plane of the cowl lip to evaluate the Mach number and angularity of the flow entering the inlet. Flighttuft stu dies. - To obtain air understanding of the complex flow around the YF- 1A ir2llet, tufts were placed on the inboard upper and lower quarters of the nacelle (ref. 24) . High-speed cameras were used to record the flow directions indicated by the tufts during supersonic flights. Tuft movement and direction were obtained along with flow patterns from the film (fig. ti). Results were obtained for a wide range of flight conditions. I'IZOGF.A11 PLANS AND SCHEDl1IYS Present Program The present Y F-12 propulsion dynamics program (table 1) includes plans to obtain compressor face and duct dynamic flight data in order to compare these results to the wind tunnel data, determine the effect of atmospheric turbulence on inlet dynamics, evaluate ca:--ses and effects of unstarts, and evaluate predictive techniques for inlet transients. The propulsion program is to be followed by the cooperative control program (table 2) . Compressor facedynamics. The objectives of these tests are to obtain dynamic compressor ace and duct pressure data in flight and to compare these data with wind tunnel data. From these comparisons dynamic distortion scaling laws can be established. For these flit,*ht tests the flexible wire that was used for the dynamic pressures is being replaced by it rigid line to obtain accurate measurements in the flight environment. Oiher propulsion instrumentation for the flight tests is being repaired during this time. The flight tests are expected to include 24 dynamic total pressure probes and 40 duct static pressure probes that have frequency responses from steady-state to 500 hertz.

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Boundary layer dynamics. - The objective of these tests is to obtain dynamic boundaryy laver pressure data to compare with the wind tunnel results. Four boundary layer rakes. each with several dynamic pressure probes, are to be positioned along the inlet duct with static pressure probes. Atmosphericall. induced turbulence. - The objective of these tests is to measure and evaluate the effects of atmospherically induced turbulence on the d y namics of mixed-compression inlets. This information is needed for ae aerodynamic design and control of inlets. The method involves the installation of .In instrument package. such as it gust probe (as used on the YF-12-935 airplane), on ti:e nose boom of the airplane to determine the upstream disturbances, such as free-stl"am turbulence and induced structural modes. The flight test is to be conducted by predicting and searching for turbulent flight conditions comparnhle to altitudes and Mach numbers of future supersonic transports. The long-range goal of this program is to develop and evaluate an instrument package that can be placed on various supersonic aircraft to obtain it large data base for the study of the effects of upstream disturbances. Inlet transients. Large-scale propulsion transients have adverse effects on aircraft stability and pasGenger comfort. A careful documentation of flight experience and a validation of prediction techni q ues tire needed for future aircraft designs. The objectives of these tests are to evaluate the causes and effects of the unstarts that occur during the compressor face and boundary layer dynamic flight test program and to compare results to analytical prediction techniques for aircraft stability and control. Cooperative control system. The overall objectives of this program are to control airplane/ propulsion system interactions, optimize total system performance. and evaluate the predicted system characteristics in flight. Strong interactions have been found to exist between engine and inlet, or between propulsion system and airframe, for supersonic aircraft, such its the XB-711. YF 12, and F 111 airplanes. Use of the cooperative control system to take advantage of favornb1c interactions and avoid or minimize unfavorable interactions could result in significant improvements in furl consumption, range, performance, and structural weight. The cooperative control program schedule is given in table 2. The program has two phases. In the first p1ase (fit*. 90a0) , the exl'<ting analog air data system, autopilot . inlet control sys'em, and autothrottle system are converted to digital conti-A. An advanced autolilot with an autothrottle has been developed, and flight tests of the complete dikital system are planned to validate the hardware and software. In the second phase (fig. 9(h)), the control systems are integrated by using control laws developed from models of the airplane's propulsion system and aerodynamics. Optimal control methods as well as classical methods lu r e beitig used to derive tho new control concepts. Flight tests are planned to validate the integrated control. A complete description of this program is given in reference 42. Expected Technical Results Propulsion system steady-state performance. - During the 8 months' aircraft down time between November 1975 and .June 1976, considerable progress is expected in the analysis of steady state data. Spec-fic areas of analysis include• steady stnt(inlet performance, steady-state inlet recovery and distortion, local inlet flc)w mess urem^nts. airflow performance and measuring techniques. and techniques for flight

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and wind tunnel comparisons of inlet rerformrxnce. Propulsion system dynamic performance. - The data reduction plan for dynam- - ic performance is shown in figure 10. The flight data, 1/3-scale data, and full scale data are to be digitized, and distortion parameters are to be calculated for selected wind tunnel/flight match points. The effects of boundary layer rakes, filters, record length, engine face rake configuration, and distortion indexes are to be investiga ted. The statistical characteristics of distortion indexes and boundary layer pressures should be evaluated and used to establish dynamic scaling laws for wind tunnel -to- flight correlations. In addition, an analysis of inlet unstarts , inlet transients. and stalls is in progress. Control systems. - The analysis of the results of the Lewis stability bleed system wind tunnel test is documented in reference 38. The flight evaluation of the T1GT control system is in progress. With the better sensing and control expected from the cooperative control system, current operational ma, gins could be reduced, resulting in payload gains. With better -Hatching of engine and inlet flow, an(! ,iith improved center of gravity control, drag can be significantly reduced. In addition, cooperative control reduces aircraft transients caused by unstart and reduces air traffic control p^oblems because of tighter flig;htpath controls. The results ohtained from this prop*.r are to be documented in NASA and contractor reports. Airframei,, —,)pulsion system interactions. - The 1/12 scale whit] tunnel pressure data are being analyzed to investigate the^c,-1 factors affecting, :iii•frame/propul sion system interactions. In addition, the 1/12-scale flow field dat;i have L ,^en reduced and are being; comparod to results obtain^d in fligTi,t . The flip*ht tuft study of thu local flow around the nacelle of the YF-12A airplane is reported in reference 24. Proposc;d. Fixture Programs Propulsion test bed. It is proposed that the YF-12 airplane be used as a propulsion test hed or evaluation of future propulsion, concepts, such as turbofan ram jets, variable cycle engine concepts, anc t supersonic transport (SST) inlets. An illustration of how the YF-12 airplane could carry test bed experiments is shown in figure 11. Test bed experiments offor several advantages. IliVh risk concepts can be investigated since the experimen t is independent of aircraft propulsion systems and has mirninal influence on airer,::'t stability and control. in addition, full-:kale models can be tested. ;return temperatures and continuous vari;it ions in Mach numher can be obtained, anti realistic flow fields can be s:nxulated. Advanced shock sensing tech nigt!es . - New shock sensing concepts are being; developed for use in control systems to operate the control, in a closed loop mode. Because of the lack of reliable sensors that function in the night environment, cur rent inlet control systems utilize a scheduled control rather than a closed loop control oxt the primary variable. The plan is to develop iiwi evaluate fiher optronics and other types of sensors in ground facilities and in flight on a ride-niong{ hasis. These sensors can lie incorporated with advanced control loops by using; the cooperative control computer.

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Shock stability bleedsystem. - The Lewis shock stability system is to be evaluate effectiveness against atmospherically induced and internal airflow disturbances. Wind tunnels are presently limited in simulating atmospherically induced disturbances to test shock stability systems. It is necessary to demonstrate that the shock stability concept is feasible in a flight environment and that mixed-compression inlets can operate nearer peak performance with such systems. This program depends on the successful accomplishment of the program to define the effects of atmospherically induced turbulence on mixed-compression inlets. Fli ht effects of annular nozzle noise suppressors. - Annular nozzles appear to be effective noise suppression devices for the SST. Forward velocity effects are needed, since most suppressor co• 2epts lose effectiveness relative to predictive static performance. The plan is astall an annular nozzle on a J58 engine to investigate reductions in flyover ^;e for high pressure ratio engines. The J58 engine is of interest because its pressure ratio is higher than that of most modern engines and it is therefore more representative of SST engines. Airplane performance and drag prediction. - Investigation of flight and wind tunnel data from the B-70 airplane program indicated several areas where additional research is required to predict the performance of vehicles of the SST type. These areas can be identified as drag polars at transonic speeds (t'ig. 12(a)), drag polars at supersonic speeds (fig. 12(b)) , lift-curve slope at high supersonic speeds (fig. 13) , and drag increment due to dumping propulsion system air overboard at off-design condition;. To investigate these subjects with the YF-12 airplane, it is necessary to obtain accurate in-flight measurements of performance, representative si.oscale models of the airplane, and accurate knowledge of the shape of the airplane in night. The Dry den Flight Research Center has the capability for most of the desired in-flight measurements, and flexibility studies have already been performed oil YF-12-935 airplane. With the addition of various subscale models, the above problem areas could be investigated so that the technology could be available for a future SST. CEN'T'ER RESEARCH AREAS The 'Y t^ -12 propulsion program is a cooperative program among the Allies, Lewis, and Dryden research centers. The nature of the program requires interdisciplinary expertise and facilities that do not exist at any single NASA center. The Dryden Flight Research Center manages and coordinates the overall program, including the YF-12 cor'racts , and performs all flight-related functions. Dryden is responsible for developing a cooperative control system for the YF-12 propulsion. system. The Lewis Research Center conducts analytical studies on inlet designs, performs fullscale wind tunnel tests, and performs engine calibration tests. Lewis is also responsible for developing new control systems that could be applicable to flight hardware. The Ames Research Centel is responsible for the analysis and design of wind tunnel models and the wind tunnel testing of these models. All three NASA centers are involved in the correlation of flight and wind tunnel data. D ► .vden Flight Research Center National AerotlULINCS and Space Adminirtration Edwards, Calif. , March 10, 1976 10

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REFERENCES Inlet Design Analysis 1 . Anderson, Bernhard H.: Design of Supersonic Inlets by a Computer Program Incorporating the Meinod of Characteristics. NASA TN U-4960, 1969. 2. Anderson, Bernhard H.: Optimization of Supersonic Inlets Using the Method of Charact:—istics. Analytical Methods in Aircraft Aerodynamics, NASA SP-228, 1970, pp. 569-581. 3. Sorensen, Virginia L.: Computer Program for Calculating Flow Fields in Supersonic Inlets. NASA TN D-2897, 1965. 4. Hingst, Warren R.; and Towne, Charles E.: Comparison of Theoretical and Experimental Boundary-Layer Development in a Mach 2.5 Mixed-Compression Inlet. NASA TM X-3026, 1974. 5. Towne, Charles E.: Evaluation of Analytical Procedures for Prediction of Turbulent Boundary Layers on a Porous Wall. NASA TM X-3063, 1974. 6. Anderson, O. L.: Finite-Difference Solution for Turbulent Swirling Compressible Flow in Axisymmetric Ducts With Strut-,. NASA CR-2365, 1974. 7. Syberg . J . ; and Hickeox , T. E.: Design of a Bleed System for a Mach 3.5 Inlet. NASA CR-2187, 1972. 8. Syberg, J.; and Koncsek, J. L.: Experimental Evaluation of a Hach 3.5 Axisymmetric Inlet. NASA CR-2563 , 1975. 9. Smeltzer., Donald B.; and Sorensen, Norman E.: Investigation of a Mixed-Compression Axisymmetric Inlet System at Mach Numbers 0.6 to 3.5. NASA TN D-6078, 1970. 10. Presley, Leroy L.: A Comparison of a Shock-Capturing Technique With Experimental Data for Thr^e-Dimensional Internal Flows. Aerodynamic Analyses Requiring Advanced Computer. , Part 1. NASA SP-347 Pt 1, 1975. 1/3-Scale Inlet Results 11. Anderson, j. Thomas; Martin. Robert K.; and Shibata, Harry H .: 1/3 Scale Inlet Model 'Pest Results. Vol. I - Test Definition and Stead y State Data Presentation . NASA CR-114702 , 1974. 12. Anderson, .1. Thomas; and Edson. Ralph D.: 1/3 Scale Inlet Model Test Results. Vol. 11 - Dynamic Data Analysis. NASA CR-114703, 1974. 13. Anderson, J. Thomas; Martin, Robert K . ; and Shibata , Harr y H.:.: 1/3 Scala Inlet Model Test Results. Vol. III - Test Definition and Steady State Da'.: Presentation. NASA CR-114704, 1974. 11

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Full-Scale Inlet Results 14. Lewis Research Center: Wind-Tunnel Installation of Full-Scale Flight Inlet of YF-12 Aircraft for Steady-State and Dynamic Evaluation. NASA TM X-3138, 1974. 15. Cubbison, Robert W. : Wind Tunnel Performance of an Isolated Full-Scale YF-12 Inlet at Mach Numbers Above 2. 1. NASA TM X-3139, 1976. 16. Cubbison , Robert W.: Effects of Angle of Attack and Flow Bypass on Wind- Tunnel Performance of an Isolated Full-Scale YF-12 Inlet at Mach Numbers Above 2.075. NASA TM X-3140, 1976. 17. Ladd, J. M.: Airflow Calibration of a J-58 Engine at Simulated Supersonic Conditions. NASA TM X-71797, 1975. Flight Test Results 18. Montoya, Earl J.: Wind-Tunnel Calibration and Requirements for In-Flight Use of Fixed Hemispherical Head Angle-of-Attack and Angle-of-Sideslip Sensors. NASA TN D-6986, 1973. 19. Johnson, Harold J . ; and Montoya, Earl J.: Local Flow Measurements at the Inlet Spike Tip of a Mach 3 Supersonic Cruise Airplane. NASA TN D-6987, 1973. 20. Burcham , Frank W. , Jr.; Montoya, Earl J . ; and Lutschg, Phillip J.: Description of YF-12C Airplane, Propulsion System, and Instrumentation for Propulsion Research Flight Tests. NASA TM X-3099, 1974. 21. Smith, R. H.; and Burcham , F. W. , Jr.: Instrumentation for In-Flight Determination of Steady-State and Dynamic Inlet Performance in Supersonic Aircraft. Instrumentation for Airbreathing Propulsion, Allen E. Fuhs and Marshall Kingery,, eds . , The MIT Press (Cambridge, Mass. ) , c. 1974, pp. 41-58. 22. Schweikhard, William G.: Test Techniques, Instrumentation, and Data Processing. Distortion Induced Engine Instability, AGARD-LS-72, Oct. 1974, pp. 6-1 -- 6-43. 23. Burcham , Frank W. , Jr.; Holzman, J,.-)n K . ; and Reukauf. , Paul J.: Preliminary Results of Flight Tests of the Propulsion System of a YF-12 Airplane at Mach Numbers to 3.0. AIAA P,,uper 73-1314, Nov. 1973. 24. Yanagidate, Craig: Tuft Study of the Local Flow Around the Nacelle of the YF-12A Airplane. NASA TM X-56035, 1975. 12

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Wind Tunnel/Flight Comparisons 25. Schweikhard , William G.; and Montoya, Earl J.: Research Instrumentation Requirements for Flight/Wind-Tunnel Tests of the YF-12 Propulsion System and Related Flight Experience. Instrumentation for Airbreathing Propulsion, Allen E. Fuhs and Marshall Kingery,, eds., The MIT Press (Cambridge, Mass.), c.1974, pp. 19-39. 26. Reukauf,, Paul J . ; Schweikhard , William G.; and Arnaiz , Henry H.: Flight-Test Techniques for Obtaining Valid Comparisons of Wind-Tunnel and Flight Results From Tests on a YF-12 Mixed-Compression Inlet. AIAA Paper 74-1195, Oct. 1974. 27. Schweikhard , William G.; and Cubbison , Robert W.: Preliminary Results From Wind Tunnel and Flight Tests of the YF- .2 Propulsion System. NASA TAI X-56016, 1973. 28. Smeltzer,, Donald B.; Smith , Ronald H.; and Cubbison , Robert W.: Wind Tunnel and Flight Performance of the YF-12 Inlet System. AIAA Paper 74-621, July 1974. Overall Propulsion System 29. Campbell. D. H.: F-12 Series Aircraft Propulsion System Performance and Development. AIAA Paper 73-821, Aug. 1973. 30. Campbell, D. H.: F-12 Inlet Development. SAE Paper 740831, Oct. 1974. Inlet and Engine Controls 31. Neiner., George H.; Arpasi , Dale J . ; and Dustin, Miles 0.: Wind-Tunnel Evaluations of YF-12 Aircraft Inlet Control System by Frequency-Response and Transient Testing. NASA TM X-3142, 1975. 32. Neiner,, George H.; Seidel , Robert C.; and Arpasi , Dale J.: Wind-Tunnel Evaluation of Experimental Controls on YF-12 Aircraft Flight Inlet by Frequency- Response and Transient Testing. NASA TM X-3143, 1975. 33. Cole, Gary L.; Cwynar, David S.; and Geyser, Lucille C.: Wind-Tunnel Evaluation of the Response of a YF-12 Aircraft Flight Inlet to Internal Airflow Perturbations by Frequency-Response Testing. NASA TM X-3141, 1974. 34. Dust-in, Miles 0.; Cole , Gary L.; and Neiner,, George H.: Continuous-Output Terminal-Shock-Position Sensor for Mixed-Compression Inlets Evaluated in Wind-Tunnel Tests of YF-12 Aircraft Inlet. NASA TM X-3144, 1974. 35. Webb, William L.; and Reukauf, Paul J.: Development of a Turbine Inlet Gas Temperature Measurement and Control System Using a Fluidic Ter..perature Sensor. AIAA Paper 73-1251, Nov. 1973. 13

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36 . Webb, W. L.: Turbine Inlet Gas Temperature Measurement and Control System. AFAPL-TR-73-116, Air Force Aero Propulsion Laboratory, Wright-Patterson Air Force Base, Dec. 1973. 37. Dustin, Miles O.; and Neiner, George II.: Evaluation by Step Response Tests of Prototype Relief Valves Designed for YF-12 Inlet Stability Bleed System. NASA TM X-3262, 1975. 38. Cole, Gary L.; Dustin, Miles O.; and Neiner, George H.: A Throat-Bypass Stability System for it YF-12 Aircraft Research Inlet Using Self-Acting Mechanical Valves. NASA 'I'M X-71779, 1975. 39. Webb, John A. , Jr.; and Dustin, Miles O.: Analysis of a Stability Valve System for Extending the Dynamic Range of a Supersonic Inlet. NASA TM X-3219, 1975. 40. Webb, W. L.; and Zewski , G. J.: J58 Cooperative Control System Study. Vols . 1, 11 and III . NASA CR- 121195 , 1973. 41. Schweikhard, William G.; and Berry, Donald T.: Cooperative Airframe/Propulsion Control for Swpersonic Cruise Aircraft. SAE Paper 740478. Apr. 1974. Reukauf , Paul J . ; Burcham . Frank W. , Jr., and Holzman, Jon K.: Status of it 42. Digital Integrated Propulsion/Flight Control System for the YF-12 Airplane. AIAA Paper 75-1180, Sept. 1975. 43. Berry. Donald T.; and Gilyard , Glenn B.: Airframe/ Propulsion Sytem Interactions --- An Important Factor in Supersonic Aircraft Flight Control. AIAA Paper 73-831, Aug. 1973. 44. Berry, 1). T.; and Gilyard , G. B.: Some Stability and Control Aspects of Air-frame/Propulsion System Interactions on the YF-12 Airplane. ASME Paper 73-WA/Aero- 4, Am. Soc . Mech . Eng . , Nov. 1973. 45. Gilyard , G. B . , Berry, 1) . T. ; and Belie, D: Analysis of a Lateral-Directional Airframe/Propulsion System Interaction of it Mach 3 Cruise Aircraft. AIAA Paper 72-961, Sept. 1972. 46. Gilyard , Glenn B.; Berry , Donald T.; and Belte , Daumants: Analysis of a Lateral-Directional Airframe/Propulsion System Interaction. NASA TM X-2829, 1973. 47. Gilyard, Glenn B.; Smith, John W.; and Falkner, Victor L.: Flight Evaluation of a Mach 3 Cruise Longitudinal Autopilot. AIAA Paper 74-910. Aug. 1974. 14

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w z 0 a n '^ rn ^ ~ ^ a t :r. z 0 a t_ a I a :a7 a _ C U c E o y u o, m m Q u E Ea C a c .1.—_ * 1 _ - _ ►2 ' 1 i U U vC G o ^ - A C g 'o a c '3 T O C ^ O C _ d x ^ y. " a v «: T C :: c. m c ^° E LhC y L t ^..^ :L F C :a. C J F o^0 ^u d L t d L Et C :a. O V :i Q 15

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n 16

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Spike bleed exit louvers Forward bypass exit louvers Aft bypass doors Translating spike Shock trap tubes `— Forward bypass doors Cowl bleed (shock trap) Spike bleed Figure 1. YF-12 inlet. 17

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Test module Airplane Facility Flight, Dryden Full-scale inlet model 10' x 10' wind tunnel, <::Z^ Lewis 113-scale inlet model 8 x 7, 9 x 7 , and 11-foot wind tunnels, Ames Full-scale engine Propuision Systems Laboratory IPSL1 Altitude Test Facility, Lewis Figure Z. Comparison of inlet configurations and facilities. 18 ItEPRODUCIpAGE: IS p(^R^ ORK3IN M

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NASA Research Scale Center E G Location of pressure sensors Type A B C I D I E I F I G Total Number of sensors Ames 113 Steady state 73 32 84 44 47 40 320 (wind tunnel I Dynamic 12 6 16 12 1 40 87 Lewis Full Steady state 75 58 128 94 45 52 402 (wind tunnel i Dynamic - 11 27 - 12 24 74 Dryden Full Steady state 9 17 33 8 13 50 3 133 (flight vehicle) Dynamic 4 4 19 6 6 24 3 66 Figu

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Figure 4. Inlet configuration showing detail of stability valve installation. REPRODUMILITY OF TIIF ORICYTNAi, PAGE IS POOH 20

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1 Engine controlg Inlet control ' Exhaust gas servos temperature rotor speed Co mnl It Pr J Mach, altitude, and attitude Aircraft rates and control accelerations servos servos i i Inlet control signals Total pressure Flight lotal temperature Stall prediction Unstart prediction t — — — — — — Int raction path without inlet control system — — — e S. Conceptual schematic d ►•awing of an integ ► •ated cont ►•ol system. Fiyu ► • Bleed Ilow pulsating / - -, - Bleed flow blowing har d and steady .,IIIIU Shock wave Separated and reverse flow regions Flow along nacelle ",^l^Jz \ It fir' i _-_7 h Vi ,gu ►-e 6. Tuft studv fo ► • high supersonic Mach number* ope? .Ution. For-%vard bypass (loot , open. 21 MRODUCIPILITY OF T11F

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Support strut -\ r Strut pivot Schlieren window-7 mpressor face -- --- - — Full-scale YF-12 inlet - Tunnel ceiling "------------- -----_---- ^- =- --Tunnel - I- Cold pipe -- I Mass flow plug Transition section Tunnel floor (u) Full--scale model. Lewis 10' X lo' runnel. Translating spike Spike tip I I Shock trap bleed tube r Slidinq cowl Porous bleed Shock trap bleed Aft bypass I ixed plug Forwa r d bypass (b) 1; 3-sc•ulf, model. Arnes Unitu►y Tunnel Facility. Figure 7. Wind funne' miler models. 22

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I I I dt ^ I I I E ^o= %v S ( I I ^^o?^ I ° > I I I I I I _y c C l,l a c A - C a I C O I L Q I '° t ^ L Uq t C z I I , C.o .E w ^ g 0V W ^ v 7 .-76 116 q oC C 4 7 W Q Ou E N /, i[° ^ A c o Y = N . C cm " ^v^? I va^E° cm, c W O I ^' J -- C I I ^ ^ I I I (U I I° °— c > > ^ v ^ SE I =^ I l a a— E I I i loiluo) luawd0janap aJPNOS uolslndoJd pup uoilp inwiS I I I I I I I I I I I I I ( I I I I ^O V O I C L I I q^ C o I I I W,u I L C t d u o G ^.r a„ o v o s I v rp O L C I r I W O O I I ^ ro^ I I $ v I ^ p+v I m / I N ,O O I p^ I o ^ I c:nao I I w^ a I I m I I I I I I I t I I I I I I I I I loiluoo aweilny 23 REPRODUCIBILITY OF THE ORIGNAL PAGE 18 P01 M

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Autopilot Air data Inlet control system Autothrottle system Propulsion system model Aerodynamic force and moment model I Digital co,.version of s:•parate systems using Fligh' test existing control laws (a) Phase 1. Derivation and mechanization of integrated control Flight test laws and real time simulation (b) Phase H. Figure 9. Cooperative cont ► •ol program. 24

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Prepare data processing equipment and software Determine effects of boundary layer rakes Determine filter effects Digitize and verify Run initial flight test case and corresponding Run check wind tunnel cases case • Flight • 113 scale • Full scale (a) Phase I. engine face Determine Determine pressure data record data Revisions 4 flight test points length reduction to software effects approach • 4 full-scale points 12 113-scale points Determine engine face rake configuration effects Initial flightlwind tunnel comparison as a function of distortion index ( b ) Phase H. Figure 10. Work flow schematic. 25

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Analyze statistical Digitize and verify remaining engine face pressure data • 22 flight test points • 5 full-scale points • 23 113-scale points (c) Phase III. characteristics of distortion indexes Statistical analysis of 200 selected pressure data Compare flight and wind tunnel data Final report Figure 10. Concluded. Figure 11. YF-12 airplane as test bed. 26 RMCIDUCIR11"'i'Y OF Ti I F ORIGINAL PAGE IS PO()K

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--0 flight-measured -- — - - Predicted based on predicted longitudinal control deflections o - Prn,: cled, based on measured longitudinal control deflections Free-stream Mach number = 1 .06 Free-stream Mach number = 1.18 .18 .16 .14 Lift 12 0 O coefficient .10 .0 r ^'J • /P I 1 11 4 1 1 11 .06 .024 .026 .028 .030 .018 .020 .022 024 .026 .028 .030 .018 .020 .022 —J Drag coefficient Drag coefficient (u) Transonic speeds. —0 flight-measured -- 'c3' - - Predicted Free -stre: m Mach number = 2.10 Free-stream Mach number = 2.50 .16 .14 .12 Lift A^ coefficient • 10 .08 .06 .018 .020 .022 .024 .010 .012 .014 016 .018 .020 .022 .024 .010 .012 .014 .016 1 1 I I I I I I i I I coefficient Drag coefficient ag (b) Supersonic speeds. Fiflure 12. Comperison of flight-measured and predicted d ►•ag. 27 RFFPRODUCIBILI':Y OF THE ORIGINAL P GL I5 POOit

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Free-stream Mach Free-stream Mach Free-stream Mach Free-stream Mach number = 1.65 number = 2.10 number = 2.50 numbe r = 1.18 .16 — .14 .12 Lift 10 coefficient .08 .06 — 0 Flight - measured — — o — — Predicted /d' i 041 2 3 4 5 6 1 2 3 4 5 6 1 2 3 4 5 6 1 2 3 4 5 6 Angle of attack , deg Angle of attack, deg Angle of attack, deg Angle of tack, deg F'iyure 13. Compa • ison between f'tiyht-measu ► •ed and predicted anyte of attack. ► 28