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- USER'S MANUAL FOR AEROFCN: A FORTRAN PROGRAM TO COMPUTE AERODYNAMIC PARAMETERS
Authors: Joseph L. Conley
Report Number: NASA-TM-104237
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This report documents the computer program AeroFcn, a utility program that computes the following 18 aerodynamic parameters: geopotential altitude, Mach number, true velocity, dynamic pressure, calibrated airspeed, equivalent airspeed, impact pressure, total pressure, total temperature, Reynolds number, speed of sound, static density, static pressure, static temperature, coefficient of dynamic viscosity, kinematic viscosity, geometric altitude, and specific energy for a standard or a modified standard day atmosphere using compressible flow and normal shock relations. Any two parameters that define a unique flight condition are selected and their values are entered interactively. The remaining parameters are computed and the solutions stored in an output file. Multiple cases can be run and the multiple case solutions can be stored in another output file for plotting. Parameter units, the output format, and primary constants in the atmospheric and
aerodynamic equations can also be changed.
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Report Date: May 1992
No. Pages: 24
Keywords: Aerodynamic equations/relations; Atmospheric equations/relations; Compressible flow; Normal shock equations/relations; Aerodynamic functions/parameters
- LINEARIZED AERODYNAMIC AND CONTROL LAW MODELS OF THE X-29A AIRPLANE AND COMPARISON WITH FLIGHT DATA
Authors: John T. Bosworth
Report Number: NASA-TM-4356
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight control system design and analysis for aircraft rely on mathematical models of the vehicle dynamics. In addition to a six-degree-of-freedom nonlinear simulation, the X-29A flight controls group developed a set of programs that calculate linear perturbation models throughout the X-29A flight envelope. The models included the aerodynamics as well as flight control system dynamics and were used for stability, controllability, and handling qualities analysis. These linear models were compared to flight test results to help provide a safe flight envelope expansion. This report presents a description of the linear models at three flight conditions and two flight control system modes. The models are presented with a level of detail that would allow the reader to reproduce the linear results if desired. Comparisons between the response of the linear model and flight-measured responses are presented to demonstrate the strengths and weaknesses of the
linear models' ability to predict flight dynamics.
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Report Date: February 1992
No. Pages: 113
Keywords: Aircraft; Flight control system; Linear aerodynamic model; State space; Rigid body dynamic models; X-29A airplane
- PRESSURE DISTRIBUTION FOR THE WING OF THE YAV-8B AIRPLANE, WITH AND WITHOUT PYLONS
Authors: Edwin J. Saltzman, John H. Del Frate, Catherine M. Sabsay and Jill M. Yarger
Report Number: NASA-TM-4429
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Pressure distribution data have been obtained in flight at four span stations on the wing panel of the YAV-8B airplane. Data obtained for the supercritical profiled wing, with and without pylons installed, ranged from Mach 0.46 to 0.88. The altitude ranged from approximately 20,000 to 40,000 ft and the resultant Reynolds numbers varied from approximately 7.2 million to 28.7 million based on the mean aerodynamic chord. Pressure distribution data and flow visualization results show that the full-scale flight wing performance is compromised because the lower surface cusp region experiences flow separation for some important transonic flight conditions. This condition is aggravated when local shocks occur on the lower surface of the wing (mostly between 20- and 35-percent chord) when the pylons are installed for Mach 0.8 and above. There is evidence that convex fairings, which cover the pylon attachment flanges, cause these local shocks. Pressure
coefficients significantly more negative than those for sonic flow also occur farther aft on the lower surface (near 60-percent chord) whether or not the pylons are installed for Mach numbers 0.8. These negative pressure coefficient peaks and associated local shocks would be expected to cause increasing wave and separation drag as transonic Mach number increases.
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Report Date: November 1992
No. Pages: 139
Keywords: Supercritical wing aerodynamics; Wing pressure distribution; Wing-pylon interaction; YAV-8B
- A PRELIMINARY LOOK AT AN OPTIMAL MULTIVARIABLE DESIGN FOR PROPULSION-ONLY FLIGHT CONTROL OF JET-TRANSPORT AIRCRAFT
Authors: Christopher P. Azzano
Report Number: NASA-CR-186014
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Control of a large jet-transport aircraft without the use of conventional control surfaces was investigated. Engine commands were used to attempt to recreate the forces and moments typically provided by the elevator, ailerons, and rudder. Necessary conditions for aircraft controllability (disturbability) were developed pertaining to aircraft configuration such as the number of engines and engine placement. An optimal linear quadratic regulator controller was developed for the Boeing 707-720,in particular,for regulation of its natural dynamic modes. The design employed a method of assigning relative weights to the natural modes, for example, phugoid and dutch roll, for a more intuitive selection of the cost function. A prototype pilot command interface was then integrated into the loop based on pseudorate command of both pitch and roll. Closed-loop dynamics were evaluated first with a batch linear simulation and then with a real-time high fidelity
piloted simulation. The NASA research pilots assisted in evaluation of closed-loop handling qualities for typical cruise and landing tasks. Recommendations for improvement on this preliminary study of optimal propulsion-only flight control are provided in the report.
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Report Date: April 1992
No. Pages: 55
Keywords: Aircraft and engine modeling; Aircraft handling qualities; Flight control systems; Modal regulator design; Propulsion-enhanced flight control
Notes: NASA Technical monitor: Glenn B. Gilyard.
- DIGITAL PERFORMANCE SIMULATION MODELS OF THE F-15, F-16XL, F-18, F-104, F-111 TACT, X-29, AND HYPERSONIC RESEARCH VEHICLE
Authors: John S. Orme
Report Number: NASA-TM-104244
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The NASA Dryden Flight Research Facility digital performance simulation program is a simplified flight simulator capable of predicting overall aircraft performance for a selected trajectory or maneuver. Seven aircraft models are available for digital performance simulation. This report describes the unique characteristics of each aircraft model and the format in which the data are stored for manipulation. The origin of aircraft-specific aerodynamic and propulsive data has been referenced whenever possible. This report also provides a quick reference and gives background information on the aircraft available for digital performance simulation.
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Report Date: January 1992
- AIRCRAFT GROUND TEST AND SUBSCALE MODEL RESULTS OF AXIAL THRUST LOSS CAUSED BY THRUST VECTORING USING TURNING VANES , Technical Memorandum
Authors: Steven A. Johnson
Report Number: NASA-TM-4341
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The NASA Dryden Flight Research Facility F/A-18 High Alpha Research Vehicle was modified to incorporate three independently controlled turning vanes located aft of the primary nozzle of each engine to vector thrust for pitch and yaw control. Ground-measured axial thrust losses were compared with the results from a 14.25 percent coldjet model for single- and dual-vanes inserted up to 25 degrees into the engine exhaust. Data are presented for nozzle pressure ratios of 2.0 and 3.0 and nozzle exit areas of 253 and 348 in2. The findings of this study indicate that subscale static nozzle test results properly predict trends but underpredict the full-scale results by approximately 1 to 4.5 percent in thrust loss.
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Report Date: January 1992
No. Pages: 29
Keywords: Axial thrust loss; Cold-jet nozzle; F/A-18 HARV; Jet exhaust turning vanes; Thrust vectoring
- VARIABLE-CAMBER SYSTEMS INTEGRATION AND OPERATIONAL PERFORMANCE OF THE AFTI/F-111 MISSION ADAPTIVE WING
Authors: John W. Smith, Wilton P. Lock and Gordon A. Payne
Report Number: NASA-TM-4370
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: he advanced fighter technology integration, the AFTI F-111 aircraft, is a preproduction F-111A testbed research airplane that was fitted with a smooth variable-camber mission adaptive wing. The camber was positioned and controlled by flexing the upper skins through rotary actuators and linkages driven by power drive units. This report describes the wing camber and control systems. The measured servoactuator frequency responses are presented along with analytical predictions derived from the integrated characteristics of the control elements. A mission adaptive wing system chronology is used to illustrate and assess the reliability and dependability of the servoactuator systems during 1524 hours of ground tests and 145 hours of flight tests.
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Report Date: April 1992
No. Pages: 43
- THRUST VECTORING FOR LATERAL-DIRECTIONAL STABILITY
Authors: Lee R. Peron and Thomas Carpenter
Report Number: NASA-CR-186016
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
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Report Date: January 1992
Notes: NASA Technical monitor: Ronald J. Ray.
- AIRBREATHING AEROSPACECRAFT FLYING QUALITY CRITERIA ASSESSMENT
Authors: Duane T. McRuer, Thomas T. Myers, Roger H. Hoh, Irving L. Ashkenas and Donald E. Johnston
Report Number: NASA-CR-4442
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A study of flying quality requirements for air-breathing aerospacecraft gives special emphasis to the unusual operational requirements and characteristics of these aircraft, including operation at hypersonic speeds. The report considers distinguishing characteristics of these vehicles, including dynamic deficiencies and their implications for control. Particular emphasis is given to the interaction of the airframe and propulsion system, and the requirements for dynamic systems integration. Past operational missions are reviewed to define tasks and maneuvers to be considered for this class of aircraft. Areas of special concern with respect to vehicle dynamics and control are identified. Experience with the space shuttle orbiter is reviewed with respect to flight control system mechanization and flight experience in approach and landing flying qualities for the National Aero-Space Plane (NASP).
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Report Date: June 1992
No. Pages: 80
Keywords: Air-breathing aerospacecraft; Approach and landing; Flying qualities
Notes: NASA Technical monitor: Mary F. Shafer.
- PRACTICAL THEORIES FOR SERVICE LIFE PREDICTION OF CRITICAL AEROSPACE STRUCTURAL COMPONENTS
Authors: William L. Ko, Richard C. Monaghan and Raymond H. Jackson
Report Number: NASA-TM-4354
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A new second-order theory was developed for predicting the service lives of aerospace structural components. The predictions based on this new theory were compared with those based on the Ko first-order theory and the classical theory of service life predictions. The new theory gives very accurate service life predictions. An equivalent constant-amplitude stress cycles method was proposed for representing the random load spectrum for crack growth calculations. This method predicts the most conservative service life. The proposed use of minimum detectable crack size, instead of proof load established crack size as an initial crack size for crack growth calculations, could give a more realistic service life.
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Report Date: March 1992
No. Pages: 27
Keywords: Fracture mechanics; Half-cycle theory; Random stress cycles; Service life prediction
Notes: Also presented as paper in proceedings of the 4th Conference on Structural Failure, Product Liability, and Technical Insurance, Vienna, Austria, July 1992. ERRATA: Figures 3, 8 and 9 (pages 17/18 and 21/22) have been corrected and substituted herein. March 1992.
- FLIGHT EVALUATION OF AN EXTENDED ENGINE LIFE MODE ON AN F-15 AIRPLANE
Authors: Lawrence P. Myers and Timothy R. Conners
Report Number: NASA-TM-104240
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An integrated flight and propulsion control system designed to reduce the rate of engine deterioration has been developed and evaluated in flight on the NASA Dryden F-15 research aircraft. The extended engine life mode increases engine pressure ratio while reducing engine airflow to lower the turbine temperature at constant thrust. The engine pressure ratio uptrim is modulated in real time based on airplane maneuver requirements, flight conditions, and engine information. The extended engine life mode logic performed well, significantly reducing turbine operating temperature. Reductions in fan turbine inlet temperature of up to 80 •F were obtained at intermediate power and up to 170 •F at maximum augmented power with no appreciable loss in thrust. A secondary benefit was the considerable reduction in thrust-specific fuel consumption. The success of the extended engine life mode is one example of the advantages gained from integrating
aircraft flight and propulsion control systems.
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Report Date: April 1992
No. Pages: 29
Keywords: Extended engine life; Reduced operating temperature at constant thrust
- POSTFLIGHT AEROTHERMODYNAMIC ANALYSIS OF PEGASUS USING COMPUTATIONAL FLUID DYNAMIC TECHNIQUES
Authors: Gary D. Kuhn
Report Number: NASA-CR-186017
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The objective of this effort was to validate the computational capability of the NASA Ames Research Center's Navier-Stokes code, F3D, for flows at high Mach numbers using comparison flight test data from the Pegasus air-launched, winged space booster. Comparisons were made with temperature and heat fluxes estimated from measurements on the wing surfaces and wing-fuselage fairing. Tests were conducted for solution convergence, sensitivity to grid density, and effects of distributing grid points to provide high density near temperature and heat-flux sensors. The measured temperatures were from sensors embedded in the ablating thermal protection system. Surface heat fluxes were from plugs fabricated of highly insulative, nonablating material, and mounted level with the surface of the surrounding ablative material. As a preflight-design tool, the F3D code produces accurate predictions of heat transfer and other aerodynamic properties, and it can provide
detailed data for assessment of boundary layer separation, shock waves, and vortex formation. As a postflight-analysis tool, the code provides a way to clarify and interpret the measured results.
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Report Date: March 1992
No. Pages: 88
Keywords: Computational fluid dynamics; Pegasus; Hypersonics; Aerothermodynamics; Wing-body aerodynamics
Notes: NASA Technical monitor: Robert E. Curry.
- EFFECTS OF COCKPIT LATERAL STICK CHARACTERISTICS ON HANDLING QUALITIES AND PILOT DYNAMICS
Authors: David G. Mitchell, Bimal L. Aponso and David H. Klyde
Report Number: NASA-CR-4443
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This report presents the results of analysis of cockpit lateral control feel-system studies. Variations in feel-system natural frequency, damping, and command sensing reference (force and position) were investigated, in combination with variations in the aircraft response characteristics. The primary data for the report were obtained from a flight investigation conducted with a variable-stability airplane, with additional information taken from other flight experiments and ground-based simulations for both airplanes and helicopters. The study consisted of analysis of handling qualities ratings and extraction of open-loop, pilot-vehicle describing functions from sum-of-sines tracking data, including, for a limited subset of these data, the development of pilot models. The study confirms the findings of other investigators that the effects on pilot opinion of cockpit feel-system dynamics are not equivalent to a comparable level of added time delay.
The effects on handling qualities ratings are, however, very similar to those of time delay, and until a more comprehensive set of criteria are developed, it is recommended that feel-system dynamics be considered a delay-inducing element in the aircraft response. The best correlation with time-delay requirements was found when the feel-system dynamics were included in the delay measurement, regardless of the command reference. This is a radical departure from past approaches.
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Report Date: June 1992
Keywords: Cockpit lateral stick characteristics; Handling qualities; Feel systems; Describing functions
Notes: NASA Technical monitor: Mary F. Shafer.
- FLIGHT TEST RESULTS OF RIBLETS AT SUPERSONIC SPEEDS
Authors: Fanny A. Zuniga, Bianca T. Anderson and Arild Bertelrud
Report Number: NASA-TM-4387
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A flight experiment to test and evaluate the skin friction drag characteristics of a riblet surface in turbulent flow at supersonic speeds was conducted at NASA Dryden. Riblets of groove sizes 0.0030 and 0.0013 in. were mounted on the F-104G flight test fixture. The test surfaces were surveyed with boundary layer rakes and pressure orifices to examine the boundary layer profiles and pressure distributions of the flow. Skin friction reductions caused by the riblet surface were reported based on measured differences of momentum thickness between the smooth and riblet surfaces obtained from the boundary layer data. Flight test results for the 0.0030 in. riblet show skin friction reductions of 4 to 8 percent for Mach numbers ranging from 1.2 to 1.6 and Reynolds numbers ranging from 2 to 3.4 million per unit foot. The results from the 0.0013 in. riblets show skin friction reductions of 4 to 15 percent for Mach 1.2 to 1.4 and Reynolds numbers ranging from
3.6 to 6 million per unit foot.
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Report Date: June 1992
No. Pages: 37
- A PRELIMINARY DESIGN AND ANALYSIS OF AN ADVANCED HEAT-REJECTION SYSTEM FOR AN EXTREME ALTITUDE ADVANCED VARIABLE CYCLE DIESEL ENGINE INSTALLED IN A HIGH-ALTITUDE ADVANCED RESEARCH PLATFORM
Authors: Richard P. Johnston
Report Number: NASA-CR-186021
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Satellite surveillance in such areas as the Antarctic indicates that from time to time concentration of ozone grows and shrinks. An effort to obtain useful atmospheric data for determining the causes of ozone depletion would require a flight vehicle capable of reaching altitudes of at least 100,000 ft and flying subsonically during the sampling portion of the mission. A study of a heat rejection system for an advanced variable cycle diesel (AVCD) engine was conducted. The engine was installed in an extreme altitude, high altitude advanced research platform. Results indicate that the waste heat from an AVCD engine propulsion system can be rejected at the maximum cruise altitude of 120,000 ft. Fifteen performance points, reflecting the behavior of the engine as the vehicle proceeded through the mission, were used to characterize the heat exchanger operation. That portion of the study is described in an appendix titled, "A Detailed Study of the Heat
Rejection System for an Extreme Altitude Atmospheric Sampling Aircraft," by a consultant, Mr. James Bourne, Lytron, Incorporated.
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Report Date: July 1992
No. Pages: 55
Keywords: Extreme altitude remotely piloted vehicle; High-altitude atmospheric sampling; Hybrid heat exchanger; Unmanned air vehicle; Variable cycle diesel
Notes: NASA Technical monitor: Russ Barber
- FLIGHT VEHICLE THERMAL TESTING WITH INFRARED LAMPS
Authors: Roger A. Fields
Report Number: NASA-TM-4336
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The verification and certification of new structural-material concepts for advanced high-speed flight vehicles relies greatly on thermal testing with infrared quartz lamps. The basic quartz heater system characteristics and design considerations are presented. Specific applications are illustrated with tests that were conducted for the X-15, the Space Shuttle, and YF-12 flight programs.
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Report Date: January 1992
No. Pages: 31
Keywords: Infrared heating; Quartz lamps; Thermostructural testing; High-temperature testing
Notes: Also presented at SEM Structural Testing Technology at High Temperatures Conference, Dayton, OH, Nov. 1991.
- FAILURE DETECTION AND FAULT MANAGEMENT TECHNIQUES FOR A PNEUMATIC HIGH-ANGLE-OF-ATTACK FLUSH AIRDATA SENSING (HI-FADS) SYSTEM
Authors: Stephen A. Whitmore, Timothy R. Moes and Cornelius T. Leondes
Report Number: NASA-TM-4335
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A high-angle-of-attack flush airdata sensing system was installed and flight tested on the F-18 High Alpha Research Vehicle at the NASA Dryden Flight Research Facility. This system uses a matrix of pressure orifices arranged in concentric circles on the nose of the vehicle to determine angles of attack, angles of sideslip, dynamic pressure, and static pressure as well as other airdata parameters of interest. Results presented use an arrangement of 11 data ports distributed symmetrically on the aircraft nose. Experience with this sensing system data indicates that the primary concern for real-time implementation is the detection and management of overall system and individual pressure sensor failures. The multiple port sensing system is more tolerant to small disturbances in the measured pressure data than conventional probe-based intrusive airdata systems. However, under adverse circumstances, large undetected failures in individual pressure ports
can result in algorithm divergence and catastrophic failure of the entire system. This paper demonstrates how system and individual port failures may be detected using x2 analysis. Once identified, the effects of failures are eliminated using weighted least squares. Background on the HI-FADS hardware, the aerodynamic model, the nonlinear regression algorithm, and x2 analysis are presented. Failure detection and fault management techniques are developed and data obtained from the High Alpha Research Vehicle flight tests are used to demonstrate the techniques.
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Report Date: January 1992
No. Pages: 22
Keywords: Airdata; Nonintrusive airdata; Fault tolerance; Pressure sensing; Chi square analysis
Notes: Also presented as AIAA 92-0263, Aerospace Sciences Meeting, Reno, NV, Jan. 1992.
- IN-FLIGHT LEADING-EDGE EXTENSION VORTEX FLOW-FIELD SURVEY MEASUREMENTS ON AN F-18 AIRCRAFT AT HIGH ANGLE OF ATTACK
Authors: David M. Richwine and David F. Fisher
Report Number: NASA-TM-4398
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flow-field measurements on the leading-edge extension (LEX) of the F-18 High Alpha Research Vehicle (HARV) were obtained using a rotating rake with 16 hemispherical-tipped five-hole probes. Detailed pressure, velocity, and flow direction data were obtained through the LEX vortex core. Data were gathered during 1-9 quasi-stabilized flight conditions at angles of attack from 10 degrees to 52 degrees and at Reynolds numbers based on mean aerodynamic chord up to 16 x 106. Normalized dynamic pressures and crossflow velocities clearly showed the primary vortex above the LEX and formation of a secondary vortex at higher angles of attack. The vortex was characterized by a ring of high dynamic pressure surrounding a region of low dynamic pressure at the vortex core center. The vortex core, subcore diameter, and vertical location of the core above the LEX increased with angle of attack. Minimum values for static pressure were obtained in the vortex subcore
and decreased nearly linearly with increasing angle of attack until vortex breakdown. Rake-measured static pressures were consistent with previously documented surface pressures and showed good agreement with flow visualization flight test results. Comparison of the LEX vortex flight test data to computational solutions at 19 degrees and 30 degrees showed fair correlation.
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Report Date: January 1992
No. Pages: 49
Notes: Also presented as AIAA Paper 91-3248 at the AIAA 9th Applied Aerodynamics Conference, Baltimore, MD, 23-25 Sep. 1991
- PILOTED SIMULATION STUDY OF A BALLOON-ASSISTED DEPLOYMENT OF AN AIRCRAFT AT HIGH ALTITUDE
Authors: James Murray, Timothy R. Moes, Ken Norlin, Jeffrey Bauer, Robert J. Geenen, Bryan J. Moulton and Stephen Hoang
Report Number: NASA-TM-104245
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A real-time piloted simulation at the NASA Dryden Flight Research Facility was used to study the feasibility of a balloon-assisted deployment of a research aircraft at high altitude. In the simulation study, an unmanned, modified sailplane was carried to 110,000 ft with a high-altitude balloon and released in a nosedown attitude. A remote pilot controlled the aircraft through a near-maximum-lift pullout and then executed a zoom climb to a trimmed, l-g flight condition. A small parachute was used to limit the Mach number during the pullout to avoid adverse transonic effects and their resultant energy losses. The use of a small rocket motor was investigated for increasing the maximum attainable altitude. Aerodynamic modifications to the basic sailplane included applying supercritical airfoil gloves over the existing wing and tail surfaces. The performance of the simulated aircraft was based on low Reynolds number wind-tunnel tests and computational
techniques. The aerodynamic model included significant Mach number and Reynolds number effects at high altitude. Parametric variations were performed to study the effects of launch altitude, gross weight, Mach number limit, and parachute size on the maximum attainable stabilized altitude. A test altitude of approximately 95,000 ft was attained, and altitudes in excess of 100,000 ft were attained with small amounts of rocketmotor assist. A discussion of the utility of this deployment technique for atmospheric science mission and other high-altitude aircraft is included.
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Report Date: January 1992
No. Pages: 33
Keywords: Atmospheric science; High-altitude aircraft; Low Reynolds number aerodynamics; Remotely piloted vehicle; Simulation
Notes: Also presented as paper for 6th Workshop on Civilian & Military Needs for Automated Instrument Platforms in the 90s and Beyond, Menlo Park, CA, Jan. 1992.
- DESIGN AND ANALYSIS ISSUES OF INTEGRATED CONTROL SYSTEMS FOR HIGH-SPEED CIVIL TRANSPORTS
Authors: Craig A. McCarty, John B. Feather, John R. Dykman, Mark A. Page and John Hodgkinson
Report Number: NASA-CR-186022
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A study was conducted to identify, rank, and define development plans for the critical guidance and control design and analysis issues as related to an economically viable and environmentally acceptable high-speed civil transport. The issues were identified in a multistep process. First, pertinent literature on supersonic cruise aircraft was reviewed, and experts were consulted to establish the fundamental characteristics and problems inherent to supersonic cruise aircraft. Next, the advanced technologies and strategies being pursued for the high-speed civil transport were considered to determine any additional unique control problems the transport may have. Finally, existing technologies and methods were examined to determine their capabilities for the design and analysis of high-speed civil transport control systems, and to identify the shortcomings and issues. Three priority levels - mandatory, highly beneficial, and desirable - were established.
Within each of these levels, the issues were further ranked. Technology development plans for each issue were defined. Each plan contains a task breakdown and schedule.
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Report Date: May 1992
No. Pages: 73
Keywords: High-speed civil transport; Integrated control systems; Technology development plans
Notes: NASA Technical monitor: Bruce Powers.
- HIGH-SPEED CIVIL TRANSPORT FLIGHT- AND PROPULSION-CONTROL TECHNOLOGICAL ISSUES
Authors: J. K. Ray, C. M. Caplin and A. A. Lambregts
Report Number: NASA-CR-186015
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This report identifies technology advances required in the flight and propulsion control system disciplines to develop a high-speed civil transport. The mission and requirements of the transport and major flight and propulsion control technology issues are discussed. Each issue is ranked and, for each issue, a plan for technology readiness is given. Certain features are unique and dominate control system design. These features include the high temperature environment, large flexible aircraft, control-configured empennage, limited flight-deck visibility, strong aerodynamic coupling between propulsion and flight control, minimizing control margins, and high availability and excellent maintainability. The failure to resolve most high-priority issues can prevent the transport from achieving its goals. The flow-time for hardware may require stimulus, since market forces may be insufficient to ensure timely production. Flight and propulsion control
technology will contribute to takeoff gross weight reduction. Similar technology advances are necessary also to ensure flight safety for the transport. The certification basis of the high-speed civil transport must be negotiated between airplane manufacturers and government regulators. Efficient, quality design of the transport will require an integrated set of design tools that support the entire engineering design team.
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Report Date: March 1992
No. Pages: 161
Keywords: High-speed civil transport
Notes: NASA Technical monitor: Bruce G. Powers.
- HIGH-SPEED CIVIL TRANSPORT ISSUES AND TECHNOLOGY PROGRAM
Authors: Marle D. Hewett
Report Number: NASA-CR-186020
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This report presents a strawman program plan consisting of technology developments and demonstrations required to support the construction of a high-speed civil transport. The plan includes a compilation of technology issues related to the development of a transport. The issues represent technical areas in which research and development are required to allow airframe manufacturers to pursue an HSCT development. The vast majority of technical issues presented require flight demonstrated and validated solutions before a transport development will be undertaken by the industry. The author believes that NASA is the agency best suited to address flight demonstration issues in a concentrated effort. The new Integrated Test Facility at NASA Dryden Flight Research Facility is considered ideally suited to the task of supporting ground validations of proof-of-concept and prototype system demonstrations before night demonstrations. An elaborate ground
hardware-in-the-loop (iron bird) simulation supported in this facility provides a viable alternative to developing an expensive full-scale prototype transport technology demonstrator. Dryden's SR-71 assets, modified appropriately, are a suitable test-bed for supporting flight demonstrations and validations of certain transport technology solutions. A subscale, manned or unmanned flight demonstrator is suitable for flight validation of transport technology solutions, if appropriate structural similarity relationships can be established. The author contends that developing a full-scale prototype transport technology demonstrator is the best alternative to ensuring that a positive decision to develop a transport is reached by the United States aerospace industry.
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Report Date: May 1992
No. Pages: 118
Keywords: High-speed civil transport; Technology development planning; Integrated Test Facility
Notes: NASA Technical monitor: Bruce G. Powers
- IN-FLIGHT EVALUATION OF AERODYNAMIC PREDICTIONS OF AN AIR-LAUNCHED SPACE BOOSTER
Authors: Robert E. Curry, Michael R. Mendenhall and Bryan Moulton
Report Number: NASA-TM-104246
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Several analytical aerodynamic design tools that were applied to the Pegasus air-launched space booster were evaluated using flight measurements. The study was limited to existing codes and was conducted with limited computational resources. The flight instrumentation was constrained to have minimal impact on the primary Pegasus missions. Where appropriate, the flight measurements were compared with computational data. Aerodynamic performance and trim data from the first two flights were correlated with predictions. Local measurements in the wing and wing-body interference region were correlated with analytical data. This complex flow region includes the effect of aerothermal heating magnification caused by the presence of a corner vortex and interaction of the wing leading edge shock and fuselage boundary layer. The operation of the first two missions indicates that the aerodynamic design approach for Pegasus was adequate, and data show that
acceptable margins were available. Additionally, the correlations provide insight into the capabilities of these analytical tools for more complex vehicles in which design margins may be more stringent.
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Report Date: April 1992
No. Pages: 21
Keywords: Aerothermodynamics; Air-launched space booster; Computational fluid dynamics; Hypersonics; Wing-body aerodynamics
Notes: Also presented at the AGARD Symposium on Theoretical and Experimental Methods in Hypersonic Flows, Torino, Italy, May 4-7, 1992.
- EFFECTS OF BLEED AIR EXTRACTION ON THRUST LEVELS OF THE F404-GE-400 TURBOFAN ENGINE
Authors: Andrew J. Yuhas and Ronald J. Ray
Report Number: NASA-TM-104247
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: ground test was performed to determine the effects of compressor bleed flow extraction on the performance of F404-GE-400 afterburning turbofan engines. The two engines were installed in the F/A-18 High Alpha Research Vehicle at the NASA Dryden Flight Research Facility. A specialized bleed ducting system was installed onto the aircraft to control and measure engine bleed airflow while the aircraft was tied down to a thrust measuring stand. The test was conducted on each engine and at various power settings. The bleed air extraction levels analyzed included flow rates above the manufacturer's maximum specification limit. The measured relationship between thrust and bleed flow extraction was shown to be essentially linear at all power settings with an increase in bleed flow causing a corresponding decrease in thrust. A comparison with the F404-GE-400 steady-state engine simulation showed the estimation to be within 1 percent of measured thrust losses
for large increases in bleed flow rate.
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Report Date: January 1992
Keywords: Bleed air extraction; Engine simulation; Forebody flow control; HARV; Pneumatic flight control; Thrust loss; Turbofan engine
Notes: Also presented as AIAA 92-3092, 28th Joint Propulsion Conference, Nashville, TN, June 1992.
- SUBSONIC FLIGHT TEST EVALUATION OF A PERFORMANCE SEEKING CONTROL ALGORITHM ON AN F-15 AIRPLANE
Authors: Glenn B. Gilyard and John S. Orme
Report Number: NASA-TM-4400
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The subsonic flight test evaluation phase of the NASA F-15 (powered by F100 engines) performance seeking control program was completed for single-engine operation at part- and military-power settings. The subsonic performance seeking control algorithm optimizes the quasi-steady-state performance of the propulsion system for three modes of operation. The minimum fuel flow mode minimizes fuel consumption. The minimum temperature mode extends turbine life by decreasing the fan turbine inlet temperature. The maximum thrust mode maximizes thrust at military power. Decreases in thrust-specific fuel consumption of 1 to 2 percent were measured in the minimum fuel flow mode; these fuel savings are significant especially for supersonic cruise aircraft. Decreases of up to approximately 100 ¡R in fan turbine inlet temperature were measured in the minimum temperature mode. Temperature reductions of this magnitude would more than double turbine life if
inlet temperature was the only life factor. Measured thrust increases of up to approximately 15 percent in the maximum thrust mode cause substantial increases in aircraft acceleration. The system dynamics of the closed-loop algorithm operation were good. The subsonic flight phase has validated the performance seeking control technology, which can significantly benefit the next generation of fighter and transport aircraft.
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Report Date: August 1992
No. Pages: 20
Keywords: Propulsion systems; Performance seeking control; Subsonic flight testing; F-15 HIDEC
Notes: Also presented as AIAA Paper 92-3743 for the 28th AIAA/SAE/ASME/ASEE Joint Propulsion Conference, Nashville, TN, July 6-8, 1992.
- SUBSONIC FLIGHT TEST EVALUATION OF A PROPULSION SYSTEM PARAMETER ESTIMATION PROCESS FOR THE F100 ENGINE
Authors: John S. Orme and Glenn B. Gilyard
Report Number: NASA-TM-4426
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Integrated engine-airframe optimal control technology may significantly improve aircraft performance. This technology requires a reliable and accurate parameter estimator to predict unmeasured variables. To develop this technology base, NASA Dryden Flight Research Facility (Edwards, CA), McDonnell Aircraft Company (St. Louis, MO), and Pratt & Whitney (West Palm Beach, FL) have developed and flight-tested an adaptive performance seeking control system which optimizes the quasi-steady-state performance of the F-15 propulsion system. This paper presents flight and ground test evaluations of the propulsion system parameter estimation process used by the performance seeking control system. The estimator consists of a compact propulsion system model and an extended Kalman filter. The extended Kalman-filter estimates five engine component deviation parameters from measured inputs. The compact model uses measurements and Kalman-filter estimates as
inputs to predict unmeasured propulsion parameters such as net propulsive force and fan stall margin. The ability to track trends and estimate absolute values of propulsion system parameters was demonstrated. For example, thrust stand results show a good correlation, especially in trends, between the performance seeking control estimated and measured thrust.
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Report Date: November 1992
No. Pages: 28
Notes: Also presented as AIAA paper, Nashville, TN.
- FLIGHT DETERMINED BENEFITS OF INTEGRATED FLIGHT-PROPULSION CONTROL SYSTEMS , Technical Memorandum
Authors: James F. Stewart, Frank W. Burcham, Jr. and Donald H. Gatlin
Report Number: NASA-TM-4393
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Over the last two decades, NASA had conducted several flight research experiments in integrated flight-propulsion control. Benefits have included improved maneuverability; increased thrust, range, and survivability; reduced fuel consumption; and reduced maintenance. The flight programs were flown at NASA Dryden Flight Research Facility, with NASA Lewis and Langley Research Centers and the Department of Defense. This paper presents the basic concepts for control integration, examples of implementation, and benefits. The F-111D experiment integrated the engine and inlet control systems. The YF-12C incorporated an integrated control system involving the inlet, autopilot, airdata, navigation, and stability augmentation systems. The F-15 research involved integration of the engine, flight, and inlet control systems. Further extension of the integration included real-time, onboard optimization of engine, inlet, and flight control variable; a
self-repairing flight control system; and an engines-only control concept for emergency control. The F-18A aircraft incorporated thrust vectoring integrated with the flight controls system to provide enhanced maneuvering at high angles of attack. The flight research programs and the resulting benefits for each are described, but particular emphasis is given to the F-15 research.
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Report Date: June 1992
No. Pages: 25
Funding Organization: WU-533-02-36
Keywords: Integrated flight-propulsion control systems; Flight controls; Propulsion
Notes: Prepared as paper 92.2.9.1 for the 18th ICAS Congress, Beijing, China, September 20-25, 1992.
- INTEGRATED FLIGHT PROPULSION CONTROL RESEARCH RESULTS USING THE NASA F-15 HIDEC FLIGHT RESEARCH FACILITY , Technical Memorandum
Authors: James F. Stewart
Report Number: NASA-TM-4394
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Over the last two decades, NASA had conducted several flight research experiments in integrated flight-propulsion control. Benefits have included improved maneuverability; increased thrust, range, and survivability; reduced fuel consumption; and reduced maintenance. The flight programs were flown at NASA Dryden Flight Research Facility. This paper presents the basic concepts for control integration, examples of implementation, and benefits of integrated flight propulsion control systems. The F-15 research involved integration of the engine, flight, and inlet control systems. Further extension of the integration included real-time, onboard optimization of engine, inlet, and flight control variables; a self-repairing flight control system; and an engines-only concept for emergency control. The flight research programs and the resulting benefits are described for the F-15 research.
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Report Date: January 1992
No. Pages: 24
Funding Organization: WU-533-02-36
Keywords: Integrated flight propulsion control; F-15 HIDEC; Performance seeking control
Notes: Prepared as AIAA paper 92-4106 for the 6th Biannual Flight Test Conference, Hilton Head, South Carolina, August 24-26, 1992.
- THERMAL-STRUCTURAL TEST FACILITIES AT NASA DRYDEN
Authors: V. Michael DeAngelis and Karl F. Anderson
Report Number: NASA-TM-104249
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The National Aero-Space Plane (NASP) has renewed interest in hypersonic flight and hot-structures technology development for both the airframe and engine. The NASA Dryden Thermostructures Research Facility is a unique national facility that has been designed to conduct thermal-mechanical tests on aircraft and aircraft components by simulating the flight thermal environment in the laboratory. The layout of the facility is presented, which includes descriptions of the high-bay test area, the instrumentation laboratories, the mechanical loading systems, and the state-of-the-art closed-loop thermal control system. The hot-structures test capability of the facility is emphasized by the Mach-3 thermal simulation conducted on the YF-12 airplane. The Liquid-Hydrogen Structural Test Facility, which is presently in the design phase, will provide the capability of thermally testing structures containing hydrogen.
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Report Date: August 1992
No. Pages: 15
Keywords: Test facilities; Thermal-structural testing
Notes: Also presented at the Society for Flight Test Engineers (SFTE) 23rd Annual Symposium, Hauppauge, NY, Aug. 3-6, 1992.
- IN-FLIGHT FLOW VISUALIZATION RESULTS FROM THE X-29 AIRCRAFT AT HIGH ANGLES OF ATTACK
Authors: John H. Del Frate and John A. Saltzman
Report Number: NASA-TM-4430
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flow visualization techniques were used on the X-29A aircraft at high angles of attack to study the vortical flow off the forebody and the surface flow on the wing and tail. The forebody vortex system was studied because asymmetries in the vortex system were suspected of inducing uncommanded yawing moments at zero sideslip. Smoke enabled visualization of the vortex system and correlation of its orientation with flight yawing moment data. Good agreement was found between vortex system asymmetries and the occurrence of yawing moments. Surface flow on the forward-swept wing of the X-29A was studied using tufts and flow cones. As angle of attack increased, separated flow initiated at the root and spread outboard encompassing the full wing by 30• angle of attack. In general, the progression of the separated flow correlated well with subscale model lift data. Surface flow on the vertical tail was also studied using tufts and flow cones. As angle of
attack increased, separated flow was observed to start at the root and spread upward. The area of separated flow on the vertical tail at angles of attack >20• correlated well with the marked decrease in aircraft directional stability.
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Report Date: November 1992
No. Pages: 17
Keywords: X-29A; Flow visualization; Forebody vortex system; Vortical flow
Notes: Also presented as AIAA Paper 92-4102 at the 6th Biennial Flight Test Conference, Hilton Head, SC, Aug. 24-26, 1992.
- FLIGHT TESTING AND SIMULATION OF AN F-15 AIRPLANE USING THROTTLES FOR FLIGHT CONTROL
Authors: Frank W. Burcham, Jr., Trindel Maine and Thomas Wolf
Report Number: NASA-TM-104255
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight tests and simulation studies using the throttles of an F-15 airplane for emergency flight control have been conducted at the NASA Dryden Flight Research Facility. The airplane and the simulation are capable of extended up-and-away flight, using only throttles for flightpath control. Initial simulation results showed that runway landings using manual throttles-only control were difficult, but possible with practice. Manual approaches flown in the airplane were much more difficult,indicating a significant discrepancy between flight and simulation. Analysis of flight data and development of improved simulation models that resolve the discrepancy are discussed. An augmented throttles-only control system that controls bank angle and flightpath with appropriate feedback parameters has also been developed, evaluated in simulations, and is planned for flight in the F-15.
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Report Date: August 1992
No. Pages: 21
Keywords: F-15; Flight test; Simulation; Propulsion-only control
Notes: Also presented as AIAA Paper 92-4109 for the 6th Biennial Flight Test Conference, Hilton Head, SC, Aug. 24, 1992.
- AEROTHERMAL TEST RESULTS FROM THE SECOND FLIGHT OF PEGASUS AIR LAUNCHED SPACE BOOSTER
Authors: Gregory A. Noffz, Timothy Moes, Edward Haering, Jr. and Paul Kolodziej,
Report Number: NASA-TM-4391
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A survey of temperature, heat-flux, and pressure measurements was obtained at speeds through Mach 8.0 on the second flight of the Pegasus air-launched space booster system. All sensors were distributed on the wing-body fairing or fillet. Sensors included thin foil-gauge thermocouples installed near the surface within the thermal protection system. Thermocouples were also installed on the surface of nonablating plugs. The resulting temperature time history allowed derivation of convective heat flux. In addition, commercially available calorimeters were installed on the fillet at selected locations. Calorimeters exhibited a larger change in measured heat flux than collocated nonablating plugs in response to particular events. Similar proportional variations in heat flux across different regions of the fillet were detected by both the calorimeters and nonablating plugs. Pressure ports were installed on some nonablating plugs to explore the effects of
port protrusion and high-frequency noise on pressure measurements. The effect of port protrusion on static-pressure measurements was found to decrease with increasing Mach number. High-frequency noise suppression was found to be desirable but not required on any future flight.
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Report Date: October 1992
No. Pages: 44
- SYSTEM OVERVIEW OF THE NASA DRYDEN INTEGRATED TEST FACILITY
Authors: Robert L. Binkley and Dale Mackall
Report Number: NASA-TM-104250
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The Integrated Test Facility, built at the NASA Dryden Flight Research Facility, provides new real-time test capabilities for emerging research aircraft. This paper is an overview of the test facility and the real-time systems developed to operate this unique facility. The facility will reduce flight test risk by minimizing the difference between the flight and ground test environments. This ground test environment is provided by combining real-time flight simulation with the actual aircraft. A brief introduction to the facility is followed by a discussion of the generic capabilities of its real-time systems. The simulation system with flight hardware and the remotely augmented vehicle system is described. An overview of many hardware systems developed for the facility follows. The benefits of applying simulation to hardware-in-the-loop testing on the X-31A flight research program concludes the paper.
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Report Date: August 1992
No. Pages: 12
Keywords: NASA Dryden; Integrated Test Facility; Systems overview; Real-time testing
Notes: Also presented at the Society of Flight Test Engineers (SFTE) 23rd Annual Symposium, Hauppage, NY, Aug. 1992.
- IN-FLIGHT SIMULATION STUDIES AT THE NASA DRYDEN FLIGHT RESEARCH FACILITY , Technical Memorandum
Authors: Mary F. Shafer
Report Number: NASA-TM-4396
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Since the late 1950’s the National Aeronautics and Space Administration’s Dryden Flight Research Facility has found in-flight simulation to be an invaluable tool. In-flight simulation has been used to address a wide variety of flying qualities questions, including low-lift-to-drag ratio approach characteristics for vehicles like the X-15, the lifting bodies, and the Space Shuttle; the effects of time delays on controllability of aircraft with digital flight-control systems, the causes and cures of pilot-induced oscillation in a variety of aircraft, and flight-control systems for such diverse aircraft as the X-15 and the X-29. In-flight simulation has also been used to anticipate problems and to avoid them and to solve problems once they appear. This paper presents an account of the in-flight simulation at the Dryden Flight Research Facility and some discussion. An extensive bibliography is included.
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Report Date: July 1992
No. Pages: 20
Funding Organization: RTOP 505-64-30
Keywords: Dryden history; Flight test; In-flight simulation
Notes: Presented as AIAA 91-2916 at the AIAA Simulation Technology Conference, New Orleans, LA, August 12–14, 1991.
- A COMPUTATIONAL FLUID DYNAMICS ANALYSIS OF THE HYPERSONIC FLIGHTS OF PEGASUS
Authors: Darren Fricker, John Mendoza and Ivan Catton
Report Number: NASA-TM-186023
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The performance of a fully viscous, three-dimensional Navier-Stokes solver, PARC3D, was tested. The criteria for judging the performance of the CFD code were based on flight data from the first two flights of the Pegasus. The flight data consisted of heat-transfer rates and sparse pressure coefficients primarily in the fillet region of the vehicle. The code performed remarkably well in all aspects of the tests. As expected, the best heat-transfer results were obtained for the low Mach number simulations. These results are attributed to lack of high ablation rates at the lower Mach numbers since the CFD simulations did not account for ablation at the vehicle surface. As the Mach number increased, the ablative effects became more apparent in the comparisons. This effect was evident at the highest Mach numbers, when PARC3D would consistently overpredict the aerothermal heat flux. To evaluate the code thoroughly, flight data would be required over the
entire fillet region, rather than at a few discrete locations. In this manner, CFD heat-flux profiles could be accurately evaluated. Although the pressure data were sparse, the trends suggest that the code predicts surface pressures with reasonable accuracy. Three of the four locations yield pressures that are consistent, but the CFD results yield pressures below the flight data by a factor of 2 at the fourth location.
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Report Date: August 1992
No. Pages: 32
Keywords: Pegasus; Computational fluid dynamics; Hypersonic flight
Notes: NASA Technical monitor: Robert Curry.
- FLIGHT-DETERMINED STABILITY ANALYSIS OF MULTIPLE-INPUT MULTIPLE-OUTPUT CONTROL SYSTEMS
Authors: John J. Burken
Report Number: NASA-TM-4416
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Singular value analysis can give conservative stability margin results. Applying structure to the uncertainty can reduce this conservatism. This paper presents flight-determined stability margins for the X-29A lateral-directional, multiloop control system. These margins are compared with the predicted unscaled singular values and scaled structured singular values. The algorithm was further evaluated with flight data by changing the roll-rate-to-aileron-command-feedback gain by ±20 percent. Minimum eigenvalues of the return difference matrix which bound the singular values are also presented. Extracting multiloop singular values from flight data and analyzing the feedback gain variations validates this technique as a measure of robustness. This analysis can be used for near-real-time flight monitoring and safety testing.
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Report Date: January 1992
Keywords: Multi-variable robustness; Singular value analysis; Structured singular values; Flight-determined singular values; Multiple-input - multiple-output stability analysis
Notes: Also presented as AIAA Paper 92-4396 at the Guidance, Navigation, and Control Conference, Hilton Head, SC, Aug. 10-12, 1992.
- THE DEVELOPMENT OF AN AIRBORNE INFORMATION MANAGEMENT SYSTEM FOR FLIGHT TEST
Authors: Glenn A. Bever
Report Number: NASA-TM-104251
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An airborne information management system is being developed at the NASA Dryden Flight Research Facility. This system will improve the state of the art in managing data acquisition on board research aircraft. The design centers around highly distributable, high-speed microprocessors that allow data compression, digital filtering, and real-time analysis. This paper describes the areas of applicability, approach to developing the system, potential for trouble areas, and reasons for this development activity. System architecture (including the salient points of what makes it unique), design philosophy, and tradeoff issues are also discussed.
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Report Date: January 1992
No. Pages: 14
Keywords: Data acquisition; Data development; Digital data; Information management; Instrumentation; Liquid cooling
Notes: Also presented as AIAA Paper 92-4113 at the 6th Biennial Flight Test Conference, Hilton Head, SC, Aug. 24-26, 1992
- THRUST STAND EVALUATION OF ENGINE PERFORMANCE IMPROVEMENT ALGORITHMS IN AN F-15 AIRPLANE
Authors: Timothy R. Conners
Report Number: NASA-TM-104252
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An investigation is underway to determine the benefits of a new propulsion system optimization algorithm in an F-15 airplane. The performance seeking control (PSC) algorithm optimizes the quasi-steady-state performance of an F100 derivative turbofan engine for several modes of operation. The PSC algorithm uses an onboard software engine model that calculates thrust, stall margin, and other unmeasured variables for use in the optimization. As part of the PSC test program, the F-15 aircraft was operated on a horizontal thrust stand. Thrust was measured with highly accurate load cells. The measured thrust was compared to onboard model estimates and to results from posttest performance programs. Thrust changes using the various PSC modes were recorded. These results were compared to benefits using the less complex highly integrated digital electronic control (HIDEC) algorithm. The PSC maximum thrust mode increased intermediate power thrust by 10
percent. The PSC engine model did very well at estimating measured thrust and closely followed the transients during optimization. Quantitative results from the evaluation of the algorithms and performance calculation models are included with emphasis on measured thrust results. The report presents a description of the PSC system and a discussion of factors affecting the accuracy of the thrust stand load measurements.
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Report Date: January 1992
Keywords: Engine performance; F100 turbofan engine; F-15 airplane; Performance seeking control; Thrust calculation; Thrust stand
Notes: Also presented as AIAA Paper 92-3747 at the 28th AIAA/SAE/ASME/ASEE Joint Propulsion Conference, Nashville, TN, July 6-8, 1992.
- F-18 SYSTEMS RESEARCH AIRCRAFT FACILITY , Technical Memorandum
Authors: Joel R. Sitz
Report Number: NASA-TM-4433
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: To help ensure that new aerospace initiatives rapidly transition to competitive U.S. technologies, NASA Dryden Flight Research Facility has dedicated a systems research aircraft facility. The primary goal is to accelerate the transition of new aerospace technologies to commercial, military, and space vehicles. Key technologies include more-electric aircraft concepts, fly-by-light systems, flush airdata systems, and advanced computer architectures. Future aircraft that will benefit are the high-speed civil transport and the National AeroSpace Plane. This paper describes the systems research aircraft flight research vehicle and outlines near-term programs.
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Report Date: December 1992
No. Pages: 36
Funding Organization: WU-505-68-52
Keywords: Fiber optics; Fly-by-light; More-electric aircraft; High-speed civil transport
Notes: Prepared as SAE paper 922063 for the Aerotech ‘92 Conference, Anaheim, California, October 5-8, 1992.
- THE F-18 HIGH ALPHA RESEARCH VEHICLE - A HIGH-ANGLE-OF-ATTACK TESTBED AIRCRAFT
Authors: Victoria A. Regenie, Donald Gatlin, Robert Kempel and Neil Matheny
Report Number: NASA-TM-104253
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
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Report Date: January 1992
Notes: Also presented as AIAA Paper 92-4121, Aug. 1992.
- OVERVIEW OF THE NASA DRYDEN FLIGHT RESEARCH FACILITY AERONAUTICAL FLIGHT PROJECTS TESTING
Authors: Robert R. Meyer, Jr.
Report Number: NASA-TM-104254
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Several principal aeronautics flight projects of the NASA Dryden Flight Research Facility are discussed. Key vehicle technology areas from a wide range of flight vehicles are highlighted. These areas include flight research data obtained for ground facility and computation correlation, applied research in areas not well suited to ground facilities (wind tunnels), and concept demonstration.
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Report Date: January 1992
Keywords: Flight test; Flow visualization; high angle of attack; Propulsion optimization flight-test techniques; Supersonic laminar flow
Notes: Prepared as AGARD 16.
- CORRELATION OF FOREBODY PRESSURES AND AIRCRAFT YAWING MOMENTS ON THE X-29A AIRCRAFT AT HIGH ANGLES OF ATTACK
Authors: David F. Fisher, David M. Richwine and Stephen Landers
Report Number: NASA-TM-4417
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: In-flight pressure distributions at four fuselage stations on the forebody of the X-29A aircraft have been reported at angles of attack from 15 to 66• and at Mach numbers from 0.22 to 0.60. At angles of attack of 20• and higher, vortices shed from the nose strake caused suction peaks in the pressure distributions that generally increased in magnitude with angle of attack. Above 30•-angle of attack, the forebody pressure distributions became asymmetrical at the most forward station, while they remained nearly symmetrical until 50- to 55•-angle of attack for the aft stations. Between 59- to 66•-angle of attack, the asymmetry of the pressure distributions changed direction. Yawing moments for the forebody alone were obtained by integrating the forebody pressure distributions. At 45¡-angle of attack, the aircraft yaws to the right and at 50• and higher, the aircraft yaws to the left. The forebody yawing moments
correlated well with the aircraft left yawing moment at an angle of attack of 50• or higher. At a 45•-angle of attack, the forebody yawing moments did not correlate well with the aircraft yawing moment, but it is suggested that this was due to asymmetric pressures on the cockpit region of the fuselage which was not instrumented. The forebody was also shown to provide a positive component of directional stability of the aircraft at angles of attack of 25• or higher. A Mach number effect was noted at angles of attack of 30• or higher at the station where the nose strake was present. At this station, the suction peaks in the presure distributions at the highest Mach number were reduced and much more symmetrical as compared to the lower Mach number pressure distributions.
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Report Date: January 1992
Keywords: Flight test; Forebody; High angle of attack; Pressure distribution; Vortex flow; X-29
Notes: Also presented as AIAA Paper 92-4105 at the Flight Test Conference, Hilton Head, SC, Aug. 24-26, 1992.
- FLIGHT TEST RESULTS FROM A SUPERCRITICAL MISSION ADAPTIVE WING WITH SMOOTH VARIABLE CAMBER , Technical Memorandum
Authors: Sheryll Goecke Powers, Lannie D. Webb, Edward L. Friend and William A. Lokos
Report Number: NASA-TM-4415
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-idle wing camber shape for any fight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 degrees, a Mach number of 0.85, leading- and trailing-edge cambers of 0/2 and 5/10, and angles of
attack from 3.0 degrees to 14.0 degrees. For the well-behaved flow of the leading- trailing-edge cambers = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the leading- trailing edge cambers = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.
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Report Date: November 1992
No. Pages: 32
Funding Organization: WU 533-02-11
Keywords: AFTI/F-111 mission adaptive wing; Boundary-layer profiles; Smooth variable-camber wing; Wing buffet intensity; Wing pressures; Wing twist
Notes: Prepared as AIAA-92-4101 for presentation at the AIAA 6th Biennial Flight Test Conference, Hilton Head, South Carolina, August 24-26, 1992.
- RAPID DEVELOPMENT OF THE X-31 SIMULATION TO SUPPORT FLIGHT TESTING
Authors: Dale Mackall, Kenneth Norlin, Dorothea Cohen, Gary Kellogg, Lawrence Schilling and John Sheen
Report Number: NASA-TM-104256
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The X-31 Enhanced Fighter Maneuverability Program has been reorganized to form the International Test Organization, with the NASA Dryden Flight Research Facility as the responsible test organization. The two X-31 research aircraft and engineering support personnel have been colocated at NASA-Dryden, with flight test operations beginning in April 1992. Therefore rapid development of a hardware-in-the-loop simulation was needed to support the flight test operations at NASA-Dryden, and to perform verification and validation of flight control software. This paper will discuss the X-31 simulation system requirements, distributed simulation system architecture, simulation components from math models to the visual system, and the advanced capabilities the X-31 simulation provides. In addition, unique software tools and the methods used to rapidly develop this simulation system will be highlighted.
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Report Date: January 1992
Keywords: Hardware-in-the-loop; Real-time simulation; Simulation cockpit; Verification and validation; Visual systems; X-31 Enhanced Fighter Maneuverability
Notes: Also presented as AIAA 92-4176, Aug. 1992.
- OVERVIEW OF SUPERSONIC LAMINAR FLOW CONTROL RESEARCH ON THE F-16XL SHIPS 1 AND 2 ,
Authors: Bianca T. Anderson and Marta Bohn-Meyer
Report Number: TM-104257
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: NASA is directing research to develop technology for a high-speed civil transport. Supersonic laminar flow control has been identified as a program element, since it offers significant drag-reduction benefits and is one of the more promising technologies for producing an economically viable aircraft design. NASA is using two prototype F-16XL aircraft to research supersonic laminar flow control. The F-16XL planform is similar to design planforms of high-speed civil transports. The planform makes the aircraft ideally suited for developing technology pertinent to high-speed transports. The supersonic laminar flow control research programs for both aircraft are described. Some general results of the ship-1 program demonstrate that significant laminar flow was obtained using laminar flow control on a highly swept wing at supersonic speeds.
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Report Date: October 1992
Keywords: Boundary layer control; F-16XL aircraft; Laminar boundary layer; Supersonic boundary layers Supersonic boundary layers
Notes: Also published as SAE paper 921994 for the Aerotech '92 Conference, Anaheim,CA, Oct. 5-8, 1992.
- A RULE-BASED SYSTEM FOR REAL-TIME ANALYSIS OF CONTROL SYSTEMS , Technical Memorandum
Authors: Richard R. Larson and D. Edward Millard
Report Number: NASA-TM-104258
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An approach to automate the real-time analysis of flight critical health monitoring and system status is being developed and evaluated at the National Aeronautics and Space Administration Dryden Flight Research Facility. A software package was developed in house and installed as part of the extended aircraft interrogation and display system. This design features a knowledge-base structure in the form of rules to formulate interpretation and decision logic of real-time data. This technique has been applied for ground verification and validation testing and flight test monitoring where quick, real-time, safety-of-flight decisions can be very critical. In many cases postprocessing and manual analysis of flight system data are not required. This paper describes the processing of real-time data for analysis and the output format, which features a message stack display. The development, construction, and testing of the rule-driven knowledge base, along
with an application using the X-31A flight test program, are presented.
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Subject Category: 31
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Report Date: October 1992
No. Pages: 15
Funding Organization: WU 505-64-30
Keywords: Control systems; Mission control center; Real-time analysis
Notes: Prepared for the IEEE/AIAA 11th Digital Avionics Systems Conference, Seattle, Washington, October 5–8, 1992.
- EXPERIENCE WITH ADA ON THE F-18 HIGH ALPHA RESEARCH VEHICLE FLIGHT TEST PROGRAM , Technical Memorandum
Authors: Victoria A. Regenie, Michael Earls, Jeanette Le and Michael Thomson
Report Number: NASA-TM-104259
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Considerable experience has been acquired with Ada at the NASA Dryden Flight Research Facility during the on-going High Alpha Technology Program. In this program, an F-18 aircraft has been highly modified by the addition of thrust-vectoring vanes to the airframe. In addition, substantial alteration was made in the original quadruplex flight control system. The result is the High Alpha Research Vehicle. An additional research flight control computer was incorporated in each of the four channels. Software for the research flight control computer was written in Ada. To date, six releases of this software have been flown. This paper provides a detailed description of the modifications to the research flight control system. Efficient ground-testing of the software was accomplished by using simulations that used the Ada for portions of their software. These simulations are also described. Modifying and transferring the Ada flight software to the software
simulation configuration has allowed evaluation of this language. This paper also discusses such significant issues in using Ada as portability, modifiability, and testability as well as documentation requirements.
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Subject Category: 05
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Report Date: October 1992
No. Pages: 17
Funding Organization: WU 533-02-35
Keywords: Ada; Airborne computers; Digital systems; F-18 High Alpha Research Vehicle; Flight control
Notes: Michael Thomson is affiliated with PRC Inc., Edwards, California. This report was also presented at the IEEE/AIAA Digital Avionics Systems Conference, October 5–8, 1992, Seattle, Washington
- THE CONSTANT CURRENT LOOP: A NEW PARADIGM FOR RESISTANCE SIGNAL CONDITIONING , Technical Memorandum
Authors: Karl F. Anderson
Report Number: NASA-TM-104260
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A practical single constant current loop circuit for the signal conditioning of variable resistance transducers has been synthesized, analyzed, and demonstrated. The strain gage and the resistance temperature device are examples of variable resistance sensors. Lead wires connect variable resistance sensors to remotely located signal conditioning hardware. The presence of lead wires in the conventional Wheatstone bridge signal conditioning circuit introduces undesired effects that reduce the quality of the data from the remote sensors. A practical approach is presented for suppressing essentially all lead wire resistance effects while indicating only the change in resistance value. Theoretical predictions supported by laboratory testing confirm the following features of the approach: (1) dc response; (2) the electrical output is unaffected by extremely large variations in the resistance of any or all lead wires; (3) the electrical output remains zero
for no change in gage resistance; (4) the electrical output is inherently linear with respect to gage resistance change; (5) the sensitivity is double that of a Wheatstone bridge circuit; and (6) the same excitation wires can serve multiple independent gages. An adaptation of current loop circuit is presented that simultaneously provides an output signal voltage directly proportional to transducer resistance change and provides temperature information that is unaffected by transducer and lead wire resistance variations. These innovations are the subject of NASA patent applications.
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Subject Category: 35
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Report Date: October 1992
No. Pages: 21
Funding Organization: WU 505-63-40
Keywords: Measuring instruments; Strain gages; Thermocouples; Test equipment; Temperature sensors
- A SUMMARY OF THE FOREBODY HIGH-ANGLE-OF-ATTACK AERODYNAMICS RESEARCH ON THE F-18 AND THE X-29A AIRCRAFT , Technical Memorandum
Authors: Lisa J. Bjarke, John H. Del Frate and David F. Fisher
Report Number: NASA-TM-104261
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: High-angle-of-attack aerodynamic studies have been conducted on both the F-18 High Alpha Research Vehicle (HARV) and the X-29A aircraft. Data obtained include on- and off-surface flow visualization and static pressure measurements on the forebody. Comparisons of similar results are made between the two aircraft where possible. The forebody shapes of the two aircraft are different and the X-29A forebody flow is affected by the addition of nose strakes and a flight test noseboom. The forebody flow field of the F-18 HARV is fairly symmetric at zero sideslip and has distinct, well-defined vortices. The X-29A forebody vortices are more diffuse and are sometimes asymmetric at zero sideslip. These asymmetries correlate with observed zero-sideslip aircraft yawing moments.
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Subject Category: 02
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Report Date: November 1992
No. Pages: 20
Funding Organization: WU 505-68-71/WU 533-02-38
Keywords: F-18 HARV; Forebody pressure distribution; High angle of attack; In-flight flow visualization; Vortex flow; X-29A aircraft
Notes: Prepared as SAE 921996 for presentation at the Society of Automotive Engineers Aerotech '92 conference, Anaheim, California, October 5–8, 1992
- STRATOSPHERIC TURBULENCE MEASUREMENTS AND MODELS FOR AEROSPACE PLANE DESIGN , Technical Memorandum
Authors: L. J. Ehernberger
Report Number: NASA-TM-104262
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Progress in computational atmospheric dynamics is exhibiting the ability of numerical simulation to describe instability processes associated with turbulence observed at altitudes between 15 and 25 km in the lower stratosphere. As these numerical simulation tools mature, they can be used to extend estimates of atmospheric perturbations from the present gust database for airplane design at altitudes below 15 km to altitudes between 25 and 50 km where aerospace plane operation would be at hypersonic speeds. The amount of available gust data and number of temperature perturbation observations are limited at altitudes between 15 and 25 km. On the other hand, in-situ gust data at higher altitudes are virtually nonexistent. The uncertain potential for future airbreathing hypersonic flight research vehicles to encounter strong turbulence at higher altitudes could penalize the design of these vehicles by undue cost or limitations on performance. Because the
atmospheric structure changes markedly with altitude, direct extrapolation of gust magnitudes and encounter probabilities to the higher flight altitudes is not advisable. This paper presents a brief review of turbulence characteristics observed in the lower stratosphere and highlights the progress of computational atmospheric dynamics that may be used to estimate the severity of atmospheric transients at higher altitudes.
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Subject Category: 47
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Report Date: December 1992
No. Pages: 28
Funding Organization: WU 763-21-51
Keywords: Clear air turbulence; Discrete gust models; Hypersonic aircraft; Mountain waves; Numerical simulations; Stratosphere
Notes: Presented at the AIAA Fourth International Aerospace Planes Conference, December 1–4, 1992, Orlando, Florida.
- HIGH-ALTITUDE BALLOON-LAUNCHED AIRCRAFT-A PILOTED SIMULATION STUDY
Authors: James E. Murray, Timothy R. Moes and Ken Norlin
Report Number: NASA-TM-104245
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A real-time piloted simulation at the NASA Dryden Flight Research Facility was used to study the feasibility of launching a research aircraft from a high-altitude balloon. In the study, the simulated aircraft was launched in a nosedown attitude at zero airspeed from 110,000 ft. After launch, the pilot flew the aircraft through a near-maximum-lift pullout and then through a zoom climb to a trimmed, 1-9 flight condition at the test altitude. The study included parametric variations to measure the effects of launch altitude, gross weight, Mach number limit, and parachute size on the test altitude attained. The aerodynamic model of the simulated aircraft was based on flight test results, low Reynolds number wind-tunnel tests, and computational models; the model included significant Mach number and Reynolds number effects at high altitude. A small parachute was included in the simulation to limit Mach number during the pull-out to avoid adverse transonic
effects and their resultant energy losses. A small rocket motor was included in the simulation and was investigated for boosting the aircraft to a higher test altitude. In the study, a test altitude of approximately 95,000 ft was attained without rocket boost, and a test altitude in excess of 100,000 ft was attained using small rocket boost.
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Report Date: January 1992
Notes: Also presented as AIAA 93-1019.
- OPTICAL AIRDATA MEASUREMENTS FOR RESEARCH FLIGHT TESTING
Authors: Rodney K. Bogue
Report Number: NASA-TM-4504
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Optical techniques for measuring airdata parameters have been demonstrated with promising results on high-performance fighter aircraft. These systems can measure the airspeed vector, and some are not as dependent on special in-flight calibration processes as current systems. Optical concepts for measuring freestream static temperature and density are feasible for in-flight applications. The best feature of these concepts is that the airdata measurements are obtained nonintrusively, and for the most part well into the freestream region of the flow field about the aircraft. Current requirements for measuring airdata at high angle of attack, and future need to measure the same information at hypersonic flight conditions place strains on existing techniques. Optical technology advances show outstanding potential for application in future programs and promise to make common use of optical concepts a reality. This paper summarizes results from several
flight-test programs and identifies the technology advances required to make optical airdata techniques practical.
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Subject Category: 06
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Report Date: January 1992
Keywords: Aerosols; Airspeed; Calibration static pressure; Doppler velocimetry; Static temperature
Notes: lso presented as AIAA Paper 92-4086 at the 6th Biennial Flight Test Conference, Hilton Head, SC, Aug. 24-26, 1992.
- COMPUTATION OF TURBULENT RECIRCULATING FLOW IN CHANNELS, AND FOR IMPINGEMENT COOLING , UCLA Ph.D dissertation
Authors: Byong Hoon Chang
Report Number: H-2289
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Fully elliptic forms of the transport equations have been solved numerically for two flow configurations. The first is turbulent flow in a channel with transverse rectangular ribs, and the second is impingement cooling of a plane surface. Both flows are relevant to proposed designs for active cooling of hypersonic vehicles using supercritical hydrogen as the coolant. Flow downstream of an abrupt pipe expansion and of a backward-facing step were also solved with various near-wall turbulence models as benchmark problems. A simple form of periodicity boundary condition was used for the channel flow with transverse rectangular ribs. The effects of various parameters on heat transfer in channel flow with transverse ribs and in impingement cooling were investigated using the Yap modified Jones and Launder low Reynold number k - e turbulence model. For the channel flow, predictions were in adequate agreement with experiment for constant property flow, with
the results for friction superior to those for heat transfer. For impingement cooling, the agreement with experiment was generally good, but the results suggest that improved modelling of the dissipation rate of turbulence kinetic energy is required in order to obtain improved heat transfer prediction, especially near the stagnation point. The k - e turbulence model was used to predict the mean flow and heat transfer for contant and variable property flows. The effect of variable properties for channel flow was investigated using the same turbulence model, but comparison with experiment yielded no clear conclusions. Also, the wall function method was modified for use in the variable properties flow with a non-adiabatic surface, and an empirical model is suggested to correctly account for the behavior of the viscous sublayer with heating.
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Report Date: June 1992
No. Pages: 123
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: This document is the result of work sponsored by NASA Dryden Flight Research Center and performed through the UCLA, Flight Systems Research Center. The contract monitor was Robert Quinn, and funding was facilitated by Dr. Kenneth Iliff.
- VARIABLE PROPERTY FLOW IN RECTANGULAR DUCTS WITH REPEATED RECTANGULAR RIB ROUGHNESS , Conference Report
Authors: B. Youn and A. F. Mills
Report Number: H-2290
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: PHOENICS was used to simulate turbulent flow of supercritical hydrogen through asymmetrically heated rectangular ducts with one side rib-roughened, in order to provide design data for active cooling systems of hypersonic aircraft where the hydrogen fuel is used as coolant. Variable property and conjugate heat transfer effects were considered. The computations using PHOENICS required provision of property sub-routines for supercritical hydrogen, a low-Reynolds number form of the k-e turbulence model, a rough-wall wall function and a procedure for calculating proper diffusion co-efficients a the fluid-solid interface. The effect of each design variable such as width and height of duct and wall thickness, on pressure drop and maximum wall temperature was investigated.
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Report Date: June 1992
No. Pages: 59
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: 4th International Phoenics User Conf., Miami, Vol 5, pp. 175-232, 1992. Document is the result of work sponsored by NASA Dryden, performed through UCLA, Flight Systems Research Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.
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