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  1. EXPERIENCE WITH ADA ON THE F-18 HIGH ALPHA RESEARCH VEHICLE FLIGHT TEST PROGRAM
    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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    Report Date: January 1993
    Keywords:      Ada; F-18 HARV; Thrust-vectoring vanes; Flight control system
    Notes: Also prepared as IEEE paper for the Digital Avionics Conference, Seattle WA, Oct. 5-8, 1993.


  2. APPLICATION OF A FLUSH AIRDATA SENSING SYSTEM TO A WING LEADING EDGE (LE-FADS)
    Authors: Stephen A. Whitmore , Timothy R. Moes , Mark W. Czerniejewski and Douglas A. Nichols
    Report Number: AIAA-93-0634
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper investigates the feasibility of locating a flush airdata sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered. The leading edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the airdata calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush airdata systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from -50• to +50• and dynamic pressures from 3.6 to 11.4 lb/ft2 were performed.
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    Subject Category: 06
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    Report Date: February 1993
    No. Pages: 14
    Keywords:      Airspeed; Angle of attack; Flush airdata; Nonintrusive airdata; Wing leading edge
    Notes: AIAA Paper for the 31st Aerospace Sciences Meeting, Reno, NV, Jan. 11-14, 1993.


  3. ENGINE EXHAUST CHARACTERISTICS EVALUATION IN SUPPORT OF ACOUSTIC TESTING , Technical Memorandum
    Authors: Kimberly A. Ennix
    Report Number: NASA-TM-104263
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden Flight Research Facility and NASA Langley Research Center completed a joint acoustic flight test program. Test objectives were (1) to quantify and evaluate subsonic climb-to-cruise noise and (2) to obtain a quality noise database for use in validating the Aircraft Noise Prediction Program. These tests were conducted using aircraft with engines that represent the high nozzle pressure ratio of future transport designs. Test flights were completed at subsonic speeds that exceeded Mach 0.3 using F-18 and F-16XL aircraft. This paper describes the efforts of NASA Dryden Flight Research Facility in this flight test program. Topics discussed include the test aircraft, setup, and matrix. In addition, the engine modeling codes and nozzle exhaust characteristics are described.
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    Subject Category: 07
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    Report Date: June 1993
    No. Pages: 14
    Funding Organization: WU 537-03-20
    Keywords:      Acoustics; Engine modeling codes; Noise; Nozzle exit velocity; Nozzle pressure ratio
    Notes: Presented at Society of Women Engineers National Conference, Chicago, Illinois, June 21–27, 1993.


  4. FLIGHT AND WIND-TUNNEL CALIBRATIONS OF A FLUSH AIRDATA SENSOR AT HIGH ANGLES OF ATTACK AND SIDESLIP AND AT SUPERSONIC MACH NUMBERS
    Authors: Timothy R. Moes, Stephen A. Whitmore and Frank L. Jordan
    Report Number: NASA-TM-104265
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A nonintrusive airdata-sensing system has been calibrated in flight and wind-tunnel experiments to an angle of attack of 70• and to angles of sideslip of ±15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, has been installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards,California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA Langley Research Center, Hampton, Virginia. The sensor consists of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and-sideslip regions studied.
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    Report Date: January 1993
    Keywords:      Airdata; Airdata calibration; FADS; Flush airdata sensor; Nonintrusive airdata
    Notes: Also presented as AIAA Paper 93-1017 for the AIAA/AHS/ASEE Aerospace Design Conference, Irvine, CA, Feb. 16-19, 1993.


  5. FLIGHT EXPERIENCE WITH LIGHTWEIGHT, LOW-POWER MINIATURIZED INSTRUMENTATION SYSTEMS
    Authors: Philip J. Hamory and James E. Murray
    Report Number: NASA-TM-4463
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Engineers at the NASA Dryden Flight Research Facility (NASA-Dryden) have conducted two flight research programs with lightweight, low-power miniaturized instrumentation systems built around commercial data loggers. One program quantified the performance of a radio-controlled model airplane. The other program was a laminar boundary-layer transition experiment on a manned sailplane. The purpose of this paper is to report NASA-Dryden personnel's flight experience with the miniaturized instrumentation systems used on these two programs. The paper will describe the data loggers, the sensors, and the hardware and software developed to complete the systems. The paper also describes how the systems were used and covers the challenges encountered to make them work. Examples of raw data and derived results will be shown as well. Finally, future plans for these systems will be discussed. For some flight research applications where miniaturized instrumentation is a requirement, the authors conclude that commercially available data loggers and sensors are viable alternatives. In fact, the data loggers and sensors make it possible to gather research-quality data in a timely and cost-effective manner.
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    Report Date: January 1993
    Keywords:      Data acquisition; Data loggers; Onboard recording; Radio-controlled model airplanes; Sailplanes; Self-contained instrumentation
    Notes: Also prepared as AIAA Paper 92-4111 for the 6th Biennial Flight Test Conference, Hilton Head, SC, Aug. 24-26, 1992. Also prepared as a journal article for the AIAA Journal of Aircraft, 1993.


  6. SUMMARY OF THE EFFECTS OF ENGINE THROTTLE RESPONSE ON AIRPLANE FORMATION FLYING QUALITIES
    Authors: Kevin R. Walsh
    Report Number: NASA-TM-4465
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight evaluation was conducted to determine the effect of engine throttle response characteristics on precision formation-flying qualities. A variable electronic throttle control system was developed and flight-tested on a TF-104G airplane with a J79-11B engine at the NASA Dryden Flight Research Facility. This airplane was chosen because of its known, very favorable thrust response characteristics. Ten research flights were flown to evaluate the effects of throttle gain, time delay, and fuel control rate limiting on engine handling qualities during a demandIng precision wing formation task. Handling quality effects of lag filters and lead compensation time delays were also evaluated. The Cooper and Harper Pilot Rating Scale was used to assign levels of handling quality. Data from pilot ratings and comments indicate that throttle control system time delays and rate limits cause significant degradations in handling qualities. Threshold values for satisfactory (level 1) and adequate (level 5) handling qualities of these key variables are presented. These results may provide engine manufacturers with guidelines to assure satisfactory handling qualities in future engine designs.
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    Report Date: January 1993
    Keywords:      Electronic engine control system; Engine throttle response; Formation-flying qualities; Handling qualities criteria; Throttle variables
    Notes: Also presented as AIAA Paper 92-3318 at the 28th Joint Propulsion Conference and Exhibit, Nashville, TN, July 6-8, 1992.


  7. OPERATIONAL AND RESEARCH ASPECTS OF A RADIO-CONTROLLED MODEL FLIGHT TEST PROGRAM , Technical Memorandum
    Authors: Gerald D. Budd, Ronald L. Gilman and David Eichstedt
    Report Number: NASA-TM-104266
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The operational and research aspects of a subscale, radio-controlled model flight test program are presented. By using low-cost free-flying models, an approach was developed for obtaining research-quality vehicle performance and aerodynamic information. The advantages and limitations learned by applying this approach to a specific flight test program are described. The research quality of the data acquired shows that model flight testing is practical for obtaining consistent and repeatable flight data.
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    Report Date: January 1993
    Keywords:      Radio control; Ducted fan engines; Free flight test apparatus
    Notes: Also presented as AIAA Paper 93-0625 for the 31st Aerospace Sciences Meeting, Reno, NV, Jan. 11-14, 1993.


  8. IN-FLIGHT INVESTIGATION OF A ROTATING CYLINDER-BASED STRUCTURAL EXCITATION SYSTEM FOR FLUTTER TESTING
    Authors: Lura Vernon
    Report Number: NASA-TM-4512
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A research excitation system was test flown at the NASA Dryden Flight Research Facility on the two-seat F-l6XL aircraft. The excitation system is a wingtip-mounted vane with a rotating slotted cylinder at the trailing edge. As the cylinder rotates during flight, the flow is alternately deflected upward and downward through the slot, resulting in a periodic lift force at twice the cylinder's rotational frequency. Flight testing was conducted to determine the excitation system's effectiveness in the subsonic, transonic, and supersonic flight regimes. Primary research objectives were to determine the system's ability to develop adequate force levels to excite the aircraft's structure and to determine the frequency range over which the system could excite structural modes of the aircraft. In addition, studies were conducted to determine optimal excitation parameters, such as sweep duration, sweep type, and energy levels. The results from the exciter were compared with results from atmospheric turbulence excitation at the same flight conditions. The comparison indicated that the vane with a rotating slotted cylinder provides superior results. The results from the forced excitation were of higher quality and had less variation than the results from atmospheric turbulence. The forced excitation data also invariably yielded higher structural damping values than those from the atmospheric turbulence data.
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    Report Date: January 1993
    Keywords:      Aeroelasticity; Flutter testing; In-flight structural excitation
    Notes: Also presented as AIAA Paper 93-1537 for the 34th Structures, Structural Dynamics, and Materials Conference, La Jolla, CA, Apr. 19-22, 1993.


  9. APPLICATION OF A FLUSH AIRDATA SENSING SYSTEM TO A WING LEADING EDGE (LE-FADS) , Technical Memorandum
    Authors: Stephen A. Whitmore, Timothy R. Moes, Mark W. Czerniejewski and Douglas A. Nichols
    Report Number: NASA-TM-104267
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper investigates the feasibility of locating a flush airdata sensing (FADS) system on a wing leading edge where the operation of the avionics or fire control radar system will not be hindered. The leading-edge FADS system (LE-FADS) was installed on an unswept symmetrical airfoil and a series of low-speed wind-tunnel tests were conducted to evaluate the performance of the system. As a result of the tests it is concluded that the aerodynamic models formulated for use on aircraft nosetips are directly applicable to wing leading edges and that the calibration process is similar. Furthermore, the agreement between the airdata calculations for angle of attack and total pressure from the LE-FADS and known wind-tunnel values suggest that wing-based flush airdata systems can be calibrated to a high degree of accuracy. Static wind-tunnel tests for angles of attack from –50 degrees to 50 degrees and dynamic pressures from 3.6 to 11.4 lb/ft(squared) were performed.
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    Subject Category: 06
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    Report Date: February 1993
    No. Pages: 17
    Funding Organization: WU 505-68-40
    Keywords:      Airspeed; Angle of attack; Flush airdata; Nonintrusive airdata; Wing leading edge
    Notes: Prepared as AIAA-93-0634 for the AIAA 31st Aerospace Sciences Meeting, Reno, Nevada, January 11–14, 1993.


  10. A PHOTOGRAMMETRIC SOLUTION TO A PARTICULAR PROBLEM
    Authors: David R. Hedgley, Jr. and Fanny A. Zuniga
    Report Number: NASA-TP-3415
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A closed-form mathematical solution to the classical photogrammetric problem is presented. Although quite general, the solution is more applicable to problems in which the image-space conjugates are very difficult to match but one of the elements of the pair is not. Additionally, observations are made that should make the solution to the general problem of automatic matching less computationally intensive. This approach was used to analyze flow visualization data for the F-18 High Alpha Research Vehicle. The conditions for this analysis were less than ideal for image-to-object-space transformation.
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    Report Date: January 1993
    Keywords:      F-18 aircraft; Flow visualization; Photogrammetry


  11. LEE WAVES: BENIGN AND MALIGNANT , Contractor Report
    Authors: M. G. Wurtele, A. Datta and R. D. Sharman
    Report Number: NASA-CR-186024
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The flow of an incompressible fluid over an obstacle will produce an oscillation in which buoyancy is the restoring force, called a gravity wave. For disturbances of this scale, the atmosphere may be treated as dynamically incompressible, even though there exists a mean static upward density gradient. Even in the linear approximation - i.e., for small disturbances - this model explains a great many of the flow phenomena observed in the lee of mountains. However, nonlinearities do arise importantly, in three ways: (i) through amplification due to the decrease of mean density with height; (ii) through the large (scaled) size of the obstacle, such as a mountain range; and (iii) from dynamically singular levels in the fluid field. These effects produce a complicated array of phenomena - Large departure of the streamlines from their equilibrium levels, high winds, generation of small scales,turbulence, etc. - that present hazards to aircraft and to lee surface areas. The nonlinear disturbances also interact with the larger-scale flow in such a manner as to impact global weather forecasts and the climatological momentum balance. If there is no dynamic barrier, these waves can penetrate vertically into the middle atmosphere (30-100 km), where recent observations show them to be of a length scale that must involve the coriolis force in any modeling. At these altitudes, the amplitude of the waves is very large, and the phenomena associated with these wave dynamics are being studied with a view to their potential impact on high performance aircraft, including the projected National Aerospace Plane (NASP). The presentation herein will show the results of analysis and of state-of-the-art numerical simulations, validated where possible by observational data, and illustrated with photographs from nature.
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    Subject Category: 47
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    Report Date: June 1993
    No. Pages: 29
    Funding Organization: WU 505-68-50, NCC 2-0374
    Keywords:      Clear air turbulence; Gravity waves; Inertia-gravity waves; Lee waves
    Notes: NASA technical monitors were K. Iliff and L.J. Ehernberger, Dryden Flight Research Facility.


  12. PROGRAM FOR AN IMPROVED HYPERSONIC TEMPERATURE-SENSING PROBE , Contractor Report
    Authors: Richard J. Reilly
    Report Number: NASA-CR-186025
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Under a NASA Dryden-sponsored contract in the mid 1960s, temperatures of up to 2200 degrees Centigrade (4000 degrees Fahrenheit) were successfully measured using a fluid oscillator. The current program, although limited in scope, explores the problem areas which must be solved if this technique is to be extended to 10,000 degrees Rankine. The potential for measuring extremely high temperatures, using fluid oscillator techniques, stems from the fact that the measuring element is the fluid itself. The containing structure of the oscillator need not be brought to equilibrium temperature with the fluid for temperature measurement, provided that a suitable calibration can be arranged. This program concentrated on review of high-temperature material developments since the original program was completed. Other areas of limited study included related pressure instrumentation requirements, dissociation, rarefied gas effects, and analysis of sensor time response.
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    Subject Category: 35
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    Report Date: June 1993
    No. Pages: 60
    Funding Organization: WU 505-68-50 NAS2-13457
    Keywords:      Temperature sensors; High temperature; Fluid oscillator
    Notes: NASA Technical Monitor: Rod Bogue, Dryden Flight Research Facility, Edwards, California.


  13. SPACE SHUTTLE HYPERSONIC AERODYNAMIC AND AEROTHERMODYNAMIC FLIGHT RESEARCH AND THE COMPARISON TO GROUND TEST RESULTS , Technical Memorandum
    Authors: Kenneth W. Iliff and Mary F. Shafer
    Report Number: NASA-TM-4499
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Aerodynamic and aerothermodynamic comparisons between flight and ground tests for the Space Shuttle at hypersonic speeds are discussed. All of the comparisons are taken from papers published by researchers active in the Space Shuttle program. The aerodynamic comparisons include stability and control derivatives, center-of-pressure location, and reaction control jet interaction. Comparisons are also discussed for various forms of heating, including catalytic, boundary layer, top centerline, side fuselage, OMS pod, wing leading edge, and shock interaction. The jet interaction and center-of-pressure location flight values exceeded not only the predictions but also the uncertainties of the predictions. Predictions were significantly exceeded for the heating caused by the vortex impingement on the OMS pods and for heating caused by the wing leading-edge shock interaction.
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    Report Date: June 1993
    No. Pages: 31
    Funding Organization: WU 506-69-50
    Keywords:      Aerothermodynamics; Boundary-layer transition; Catalytic effects; Flight-to-ground correlation; Hypersonic aerodynamics
    Notes: Presented as AIAA 92-3988 at the 17th Aerospace Ground Testing Conference, Nashville, Tennessee, July 6-8, 1992.


  14. COMPRESSIVE AND SHEAR BUCKLING ANALYSIS OF METAL MATRIX COMPOSITE SANDWICH PANELS UNDER DIFFERENT THERMAL ENVIRONMENTS , Technical Memorandum
    Authors: William L. Ko
    Report Number: NASA-TM-4492
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Combined inplane compressive and shear buckling analysis was conducted on flat rectangular sandwich panels using the Rayleigh-Ritz minimum energy method with a consideration of transverse shear effect of the sandwich core. The sandwich panels were fabricated with titanium honeycomb core and laminated metal matrix composite face sheets. The results show that slightly slender (along unidirectional compressive loading axis) rectangular sandwich panels have the most desirable stiffness-to-weight ratios for aerospace structural applications; the degradation of buckling strength of sandwich panels with rising temperature is faster in shear than in compression; and the fiber orientation of the face sheets for optimum combined-load buckling strength of sandwich panels is a strong function of both loading condition and panel aspect ratio. Under the same specific weight and panel aspect ratio, a sandwich panel with metal matrix composite face sheets has much higher buckling strength than one having monolithic face sheets.
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    Report Date: June 1993
    No. Pages: 23
    Keywords:      Buckling interaction surfaces; Combined load buckling; Metal matrix composites; Sandwich panels
    Notes: Also presented at the 7th International Conference on Composite Structures (ICCS), Paisley, Scotland, July 1993.


  15. ACTUATOR AND AERODYNAMIC MODELING FOR HIGH-ANGLE-OF-ATTACK AEROSERVOELASTICITY , Technical Memorandum
    Authors: Martin J. Brenner
    Report Number: NASA-TM-4493
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Accurate prediction of airframe/actuation coupling is required by the imposing demands of modern flight control systems. In particular, for agility enhancement at high angle of attack and low dynamic pressure, structural integration characteristics such as hinge moments, effective actuator stiffness, and airframe/control surface damping can have a significant effect on stability predictions. Actuator responses are customarily represented with low-order transfer functions matched to actuator test data, and control surface stiffness if often modeled as a linear spring. The inclusion of the physical properties of actuation and its installation on the airframe is therefore addressed in this paper using detailed actuator models which consider the physical, electrical, and mechanical elements of actuation. The aeroservoelastic analysis procedure is described in which the actuators are modeled as detailed high-order transfer functions and as approximate low-order transfer functions. The impacts of unsteady aerodynamic modeling on aeroservoelastic stability are also investigated in this paper by varying the order of approximation, or number of aerodynamic lag states, in the analysis. Test data from a thrust-vectoring configuration of an F/A-18 aircraft are compared to predictions to determine the effects on accuracy as a function of modeling complexity.
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    Report Date: January 1993
    Keywords:      Actuator dynamics; Aeroservoelasticity; High angle of attack; Modal stability; Structural dynamics; Unsteady aerodynamics
    Notes: Also presented as AIAA Paper 93-1419 at the AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, La Jolla, CA, Apr. 19-22, 1993.


  16. GROUND VIBRATION AND FLIGHT FLUTTER TESTS OF THE SINGLE-SEAT F-16XL AIRCRAFT WITH A MODIFIED WING
    Authors: David F. Voracek
    Report Number: NASA-TM-104264
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA single-seat F-16XL aircraft was modified by the addition of a glove to the left wing. Vibration tests were conducted on the ground to assess the changes to the aircraft caused by the glove. Flight flutter testing was conducted on the aircraft with the glove installed to ensure that the flight envelope was free of aeroelastic or aeroservoelastic instabilities. The ground vibration tests showed that above 20 Hz, several modes that involved the control surfaces were significantly changed. Flight test data showed that modal damping levels and trends were satisfactory where obtainable. The data presented in this report include estimated modal parameters from the ground vibration and flight flutter test.
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    Report Date: January 1993
    Keywords:      Aeroelasticity; Flight flutter testing; Ground vibration testing


  17. ON THE ESTIMATION ALGORITHM USED IN ADAPTIVE PERFORMANCE OPTIMIZATION OF TURBOFAN ENGINES , Technical Memorandum
    Authors: Martin D. Espana and Glenn B. Gilyard
    Report Number: NASA-TM-4551
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The performance seeking control algorithm is designed to continuously optimize the performance of propulsion systems. The performance seeking control algorithm uses a nominal model of the propulsion system and estimates, in flight, the engine deviation parameters characterizing the engine deviations with respect to nominal conditions. In practice, because of measurement biases and/or model uncertainties, the estimated engine deviation parameters may not reflect the engine’s actual off-nominal condition. This factor has a necessary impact on the overall performance seeking control scheme exacerbated by the open-loop character of the algorithm. In this report, the effects produced by unknown measurement biases over the estimation algorithm are evaluated. This evaluation allows for identification of the most critical measurements for application of the performance seeking control algorithm to an F100 engine. An equivalence relation between the biases and engine deviation parameters stems from an observability study; therefore, it is undecided whether the estimated engine deviation parameters represent the actual engine deviation or whether they simply reflect the measurement biases. A new algorithm, based on the engine’s (steady-state) optimization model, is proposed and tested with flight data. When compared with previous Kalman filter schemes, based on local engine dynamic models, the new algorithm is easier to design and tune and it reduces the computational burden of the onboard computer.
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    Subject Category: 07
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    Report Date: December 1993
    No. Pages: 36
    Funding Organization: RTOP 533-02-39
    Keywords:      Adaptive optimization; Measurement biases influence; Parameter estimation; Performance seeking control; Propulsion systems
    Notes: Prepared as paper 93-1823 for the AIAA Joint Propulsion Conference, June 28-July 1, 1993, Monterey, California.


  18. PRELIMINARY SUPERSONIC FLIGHT TEST EVALUATION OF PERFORMANCE SEEKING CONTROL
    Authors: John S. Orme and Glenn B. Gilyard
    Report Number: NASA-TM-4494
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Digital flight and engine control, powerful onboard computers, and sophisticated controls techniques may improve aircraft performance by maximizing fuel efficiency, maximizing thrust, and extending engine life. An adaptive performance seeking control system for optimising the quasi-steady state performance of an F-15 aircraft has been developed and flight tested. This system has three optimization modes: minimum fuel, maximum thrust, and minimum fan turbine inlet temperature. Tests of the minimum fuel and fan turbine inlet temperature modes were performed at a constant thrust. Supersonic single-engine flight tests of the three modes were conducted using varied afterburning power settings. At supersonic conditions, the performance seeking control law optimizes the integrated airframe, inlet, and engine. At subsonic conditions, only the engine is optimized. Supersonic flight tests showed improvements in thrust of 9 percent, increases in fuel savings of 8 percent, and reductions of up to 85 •R in turbine temperatures for all three modes. This paper describes the supersonic performance seeking control structure and gives preliminary results of supersonic performance seeking control tests. These findings have implications for improving performance of civilian and military aircraft.
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    Report Date: January 1993
    Keywords:      Afterburner controls; Aircraft flight tests; Aircraft performance; Engine control systems; Inlet controls; Propulsion systems
    Notes: Also prepared as AIAA Paper 93-1821 for the AIAA/SAE/ASME/ASEE Joint Propulsion Conference, June 28-30, 1993


  19. FLIGHT-DETERMINED ENGINE EXHAUST CHARACTERISTICS OF AN F404 ENGINE IN AN F-18 AIRPLANE , Technical Memorandum
    Authors: Kimberly A. Ennix, Frank W. Burcham, Jr. and Lannie D.Webb
    Report Number: NASA-TM-4538
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Personnel at the NASA Langley Research Center (NASA-Langley) and the NASA Dryden Flight Research Facility (NASA-Dryden) have recently completed a joint acoustic flight test program. Several types of aircraft with high nozzle pressure ratio engines were flown to satisfy a twofold objective. First, assessments were made of subsonic climb-to-cruise noise from flights conducted at varying altitudes in a Mach 0.30 to 0.90 range. Second, using data from flights conducted at constant altitude in a Mach 0.30 to 0.95 range, engineers obtained a high-quality noise database. This database was desired to validate the Aircraft Noise Prediction Program and other system noise prediction codes. NASA-Dryden personnel analyzed the engine data from several aircraft that were flown in the test program to determine the exhaust characteristics. The analysis of the exhaust characteristics from the F-18 aircraft will be reported in this paper. This paper presents an overview of the flight test planning, instrumentation, test procedures, data analysis, engine modeling codes, and results.
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    Subject Category: 07
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    Report Date: October 1993
    No. Pages: 14
    Funding Organization: WU-537-03-20
    Keywords:      Acoustics; Climb to cruise; Engine exhaust characteristics; Environmental impact
    Notes: Also prepared as AIAA Paper 93-2543 for the AIAA/SAE/ASME/ASEE 29th Joint Propulsion Conference and Exhibit, Monterey, CA, June 28-30, 1993.


  20. PRELIMINARY FLIGHT RESULTS OF A FLY-BY-THROTTLE EMERGENCY FLIGHT CONTROL SYSTEM ON AN F-15 AIRPLANE , Technical Memorandum
    Authors: Frank W. Burcham, Jr., Trindel A. Maine, C. Gordon Fullerton and Edward A. Wells
    Report Number: NASA-TM-4503
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A multi-engine aircraft, with some or all of the flight control system inoperative, may use engine thrust for control. NASA Dryden has conducted a study of the capability and techniques for this emergency flight control method for the F-15 airplane. With an augmented control system, engine thrust, along with appropriate feedback parameters, is used to control flightpath and bank angle. Extensive simulation studies have been followed by flight tests. This paper discusses the principles of throttle-only control, the F-15 airplane, the augmented system, and the flight results including actual landings with throttle-only control.
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    Subject Category: 08
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    Report Date: June 1993
    No. Pages: 19
    Funding Organization: RTOP 533-02-34
    Keywords:      F-15 airplane; Flight control; Flight-propulsion control integration; Propulsion control; Throttles-only control
    Notes: Also presented as AIAA Paper 93-1820 at the 29th AIAA/SAE/ASME Joint Propulsion Conference, Monterey, CA, June 28-30, 1993.


  21. EXTRACTION OF STABILITY AND CONTROL DERIVATIVES FROM ORBITER FLIGHT DATA , Technical Memorandum
    Authors: Kenneth W. Iliff and Mary F. Shafer
    Report Number: NASA-TM-4500
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The Space Shutter Orbiter has provided unique and important information on aircraft flight dynamics. This information has provided the opportunity to assess the flight-derived stability and control derivatives for maneuvering flight in the hypersonic regime. In the case of the Space Shuttle Orbiter, these derivatives are required to determine if certain configuration placards (limitations on the flight envelope) can be modified. These placards were determined on the basis of preflight predictions and the associated uncertainties. As flight-determined derivatives are obtained, the placards are reassessed, and some of them are removed or modified. Extraction of the stability and control derivatives was justified by operational considerations and not by research consideration. Using flight results to update the predicted database of the orbiter is one of the most completely documented processes for a flight vehicle. This process followed from the requirement for analysis of flight data for control system updates and for expansion of the operational flight envelope. These results show significant changes in many important stability and control derivatives from the preflight database. This paper presents some of the stability and control derivative results obtained from Space Shuttle flights. Some of the limitations of this information are also examined.
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    Subject Category: 02
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    Report Date: June 1993
    No. Pages: 50
    Funding Organization: WU 505-68-50
    Keywords:      Aerodynamics; Flight-to-wind-tunnel comparison; Hypersonics; Space Shuttle; Stability and control derivatives
    Notes: Previously presented at the Orbiter Experiments (OEX) Aerothermodynamics Symposium, April 27-30, 1993, Williamsburg, Virginia.


  22. FLIGHT VALIDATION OF A PULSED SMOKE FLOW VISUALIZATION SYSTEM
    Authors: Donald T. Ward and Kenneth M. Dorsett
    Report Number: NASA-CR-186026
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flow visualization scheme, designed to measure vortex fluid dynamics on research aircraft, was validated in flight. Strake vortex trajectories and axial core velocities were determined using pulsed smoke, high-speed video images, and semiautomated image edge detection hardware and software. Smoke was pulsed by using a fast-acting three-way valve. After being redesigned because of repeatedly jamming in flight, the valve shuttle operated flawlessly during the last two tests. A 25-percent scale, Gothic strake was used to generate vortex over the wing of a GA-7 Cougar and was operated at a local angle of attack of 22¡ and Reynolds number of approximately 7.8 x 10 to the 5th power/ft. Maximum axial velocities measured in the vortex core were between 1.75 and 1.95 times the freestream velocity. Analysis of the pulsed smoke system's effect on forebody vortices indicates that the system may reorient the forebody vortex system; however, blowing momentum coefficients normally used will have no appreciable effect on the leading-edge extension vortex system. It is recommended that a similar pulsed smoke system be installed on the F/A-18 High Alpha Research Vehicle and that this approach be used to analyze vortex core dynamics during the remainder of its high-angle-of-attack research flights.
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    Report Date: January 1993
    Keywords:      Flight test; Flow visualization; Smoke pulsation; Vortex fluid dynamics


  23. RECENT FLIGHT-TEST RESULTS OF OPTICAL AIRDATA TECHNIQUES , Technical Memorandum
    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: May 1993
    No. Pages: 19
    Funding Organization: RTOP 505-68-50
    Keywords:      Aerosols; Airspeed; Calibration static pressure; Doppler velocimetry; Static temperature
    Notes: Presented as SAE921997 at the Society of Automotive Engineers Aerotech ‘92, Anaheim, California, October 5-8, 1992.


  24. PERSPECTIVE ON THE NATIONAL AERO-SPACE PLANE PROGRAM INSTRUMENTATION DEVELOPMENT , NASA TM
    Authors: Rodney K. Bogue and Peter Erbland
    Report Number: NASA-TM-4505
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents a review of the requirement for, and development of, advanced measurement technology for the National Aero-Space Plane program. The objective is to discuss the technical need and the program commitment required to ensure that adequate and timely measurement capabilities are provided for ground and flight testing in the NASP program. The paper presents the scope of the measurement problem, describes the measurement process, examines how instrumentation technology development has been affected by NASP program evolution, discusses the national effort to define measurement requirements and assess the adequacy of current technology to support the NASP program and summarizes the measurement requirements. The unique features of the NASP program that complicate the understanding of requirements and the development of viable solutions are illustrated.
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    Subject Category: 06
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    Report Date: May 1993
    No. Pages: 18
    Funding Organization: 763-21-41 RX
    Keywords:      Aircraft Combustion; Heat transfer; High temperature instrumentation; Laser-induced fluorescence; National Aero-Space Plan; Scramjet; Strain
    Notes: Prepared as AIAA paper at the Aero Space Planes Conference, Orlando, FL, Dec. 1–3, 1992.


  25. PRELIMINARY DESIGN OF AN INTERMITTENT SMOKE FLOW VISUALIZATION SYSTEM
    Authors: Donald T. Ward and James H. Myatt
    Report Number: NASA-CR-186027
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A prototype intermittent flow visualization system has been designed to study vortex flow field dynamics has been constructed and tested through its ground test phase. It produces discrete pulses of dense white smoke consisting of particles of terephthalic acid by the pulsing action of a fast-acting three-way valve. The trajectories of the smoke pulses can be tracked by a video imaging system without intruding in the flow around in flight. Two methods of pulsing the smoke were examined. The simplest and safest approach is to simply divert the smoke between the two outlet ports on the valve; this approach should be particularly effective if it were desired to inject smoke at two locations during the same test event. The second approach involves closing off one of the outlet ports to momentarily block the flow. The second approach requires careful control of valve dwell times to avoid excessive pressure buildup within the cartridge container and does also increase the velocity of the smoke injected into the flow. The flow of the smoke has been blocked for periods ranging from 30 to 80 milliseconds, depending on the system volume and the length of time the valve is allowed to remain open between valve closings.
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    Keywords:      Flow visualization; Smoke generation; Smoke grenades; Smoke pulsation


  26. TEACHING HIGH-PERFORMANCE SKILLS USING ABOVE-REAL-TIME TRAINING
    Authors: Dutch Guckenberger, Kevin C. Uliano and Norman E. Lane
    Report Number: NASA-CR-4528
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The above real-time training (ARTT) concept is an approach to teaching high-performance skills. ARTT refers to a training paradigm that places the operator in a simulated environment that functions at faster than normal time. It represents a departure from the intuitive, but not often supported, feeling that the best practice is determined by the training environment with the highest fidelity. This approach is hypothesized to provide greater "transfer value" per simulation trial, by incorporating novel training techniques and instructional features into the simulator. This report discusses two related experiments. In the first, 25 naive male subjects performed three tank gunnery tasks on a simulator under varying levels of time acceleration (i.e., 1.0x, 1.6x, 2.0x, sequential, and mixed). They were then transferred to a standard (1.0x) condition for testing. Every accelerated condition or combination of conditions produced better training and transfer than the standard condition. Most effective was the presentation of trails at 1.0x, 1.6x, and 2.0x in a random order during training. Overall, the best ARTT group scored about 50 percent higher and trained in 25 percent less time compared to the real-time control group. In the second experiment, 24 mission-capable tasks on a part-task F16A flight simulator under varying levels of time compression (i.e., 1.0x, 1.5x, 2.0x, and random). All subjects were then tested in a real-time environment. The emergency procedure (EP) task results showed increased accuracy for the ARTT groups. In testing (transfer), the ARTT groups not only performed the EP more accurately, but dealt with a simultaneous enemy significantly better than a real-time control group. Although the findings on an air combat maneuvering task and stern conversion task were mixed, most measures indicated that the ARTT groups performed better and faster than a real-time control group. Other implications for ARTT are discussed along with future research directions.
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    Report Date: January 1993
    Keywords:      Above-real-time training; Gunnery simulation; High-performance skills
    Notes: NASA Technical monitor: Jack Kolf.


  27. PERFORMANCE-SEEKING CONTROL: PROGRAM OVERVIEW AND FUTURE DIRECTIONS , Technical Memorandum
    Authors: Glenn B. Gilyard and John S. Orme
    Report Number: NASA-TM-4531
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight test evaluation of the performance-seeking control (PSC) algorithm on the NASA F-15 highly integrated digital electronic control research aircraft was conducted for single-engine operation at subsonic and supersonic speeds. The model-based PSC system was developed with three optimization modes: mininum fuel flow at constant thrust, minimum turbine temperature at constant thrust, and maximum thrust at maximum dry and full afterburner throttle settings. Subsonic and supersonic flight testing were conducted at the NASA Dryden Flight Research Facility covering the three PSC optimization modes and over the full throttle range. Flight results show substantial benefits. In the maximum thrust mode, thrust increased up to 15 percent at subsonic and 10 percent at supersonic flight conditions. The minimum fan turbine inlet temperature mode reduced temperatures by more than 100 •F at high altitudes. The minimum fuel flow mode results decreased fuel consumption up to 2 percent in the subsonic regime and almost 10 percent supersonically. These results demonstrate that PSC technology can benefit the next generation of fighter or transport aircraft. NASA Dryden is developing an adaptive aircraft performance technology system that is measurement based and uses feedback to ensure optimality. This program will address the technical weaknesses identified in the PSC program and will increase performance gains.
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    Subject Category: 07
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    Report Date: August 1993
    No. Pages: 19
    Funding Organization: WU 533-02-39
    Keywords:      F-15 HIDEC; Performance optimization; Performance seeking control; Propulsion systems; Subsonic flight testing.
    Notes: Prepared as AIAA Paper 93-3765 for the Guidance, Navigation, and Control Conference, Monterey, CA, Aug. 9-11


  28. THE DEVELOPMENT AND FLIGHT TEST OF A DEPLOYABLE PRECISION LANDING SYSTEM FOR SPACECRAFT RECOVERY , Technical Memorandum
    Authors: Alex G. Sim, James E. Murray, David C. Neufeld and R. Dale Reed
    Report Number: NASA-TM-4525
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A joint NASA Dryden Flight Research Facility and Johnson Space Center program was conducted to determine the feasibility of the autonomous recovery of a spacecraft using a ram-air parafoil system for the final stages of entry from space that included a precision landing. The feasibility of this system was studied using a flight model of spacecraft in the generic shape of a flattened biconic which weighed approximately 150 lb and was flown under a commercially available, ram-air parachute. Key elements of the vehicle included the Global Positioning System guidance for navigation, flight control computer, ultrasonic sensing for terminal altitude, electronic compass, and onboard data recording. A flight test program was used to develop and refine the vehicle. This vehicle completed an autonomous flight from an altitude of 10,000 ft and a lateral offset of 1.7 miles which resulted in a precision flare and landing into the wind at a predetermined location. At times, the autonomous flight was conducted in the presence of winds approximately equal to vehicle airspeed. Several novel techniques for computing the winds postflight were evaluated. Future program objectives are also presented.
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    Subject Category: 16
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    Report Date: September 1993
    No. Pages: 33
    Funding Organization: WU 505-68-50
    Keywords:      Autonomous landing systems; Global positioning system navigation; Parachute decelerator; Ram-air parachutes; Satellite-based air navigation; Wind estimation
    Notes: R. Dale Reed is affiliated with PRC Inc., Edwards, California.


  29. FLIGHT TESTING OF AIRBREATHING HYPERSONIC VEHICLES , Technical Memorandum
    Authors: John W. Hicks
    Report Number: NASA-TM-4524
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Using the scramjet engine as the prime example of a hypersonic airbreathing concept, this paper reviews the history of and addresses the need for hypersonic flight tests. It also describes how such tests can contribute to the development of airbreathing technology. Aspects of captive-carry and free-flight concepts are compared. An incremental flight envelope expansion technique for manned flight vehicles is also described. Such critical issues as required instrumentation technology and proper scaling of experimental devices are addressed. Lastly, examples of international flight test approaches, existing programs, or concepts currently under study, development, or both, are given.
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    Subject Category: 18
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    Report Date: October 1993
    No. Pages: 39
    Funding Organization: WU 505-70
    Keywords:      Airbreathing hypersonic vehicles; Flight tests; Ground tests; Spacecraft design; Spacecraft development.
    Notes: Presented as paper no. 37 at Space Course 1993, October 11-12, 1993, Munich, Germany.


  30. REDUCTION OF STRUCTURAL LOADS USING MANEUVER LOAD CONTROL ON THE ADVANCED FIGHTER TECHNOLOGY INTEGRATION (AFTI)/F-111 MISSION ADAPTIVE WING , Technical Memorandum
    Authors: Stephen V. Thornton
    Report Number: NASA-TM-4526
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A transonic fighter-bomber aircraft, having a swept supercitical wing with smooth variable-camber flaps was fitted with a maneuver load control (MLC) system that implements a technique to reduce the inboard bending moments in the wing by shifting the spanwise load distribution inboard as load factor increases. The technique modifies the spanwise camber distribution by automatically commanding flap position as a function of flap position, true airspeed, Mach number, dynamic pressure, normal acceleration, and wing sweep position. Flight test structural loads data were obtained for loads in both the wing box and the wing root. Data from uniformly deflected flaps were compared with data from flaps in the MLC configuration where the outboard segment of three flap segments was deflected downward less than the two inboard segments. The changes in the shear loads in the forward wing spar and at the roots of the stabilators also are presented. The camber control system automatically reconfigures the flap through varied flight conditions. Configurations having both moderate and full trailing-edge flap deflection were tested. Flight test data were collected at Mach numbers of 0.6, 0.7, 0.8, and 0.9 and dynamic pressures of 300, 400, 600, and 800 lb/ft2. The Reynolds numbers for these flight conditions ranged from 26 x 10 to the 6th power to 54 x 10 to the 6th power at the mean aerodynamic chord. Load factor increases of up to 1.0 g achieved with no increase in wing root bending moment with the MLC flap configuration.
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    Subject Category: 02
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    Report Date: September 1993
    No. Pages: 32
    Funding Organization: WU 505-63-50
    Keywords:      Load alleviation; Load distribution; Maneuver load control; Variable camber; Wing bending moment


  31. CORRELATION OF ANALYTICAL AND EXPERIMENTAL HOT STRUCTURE VIBRATION RESULTS
    Authors: Michael W. Kehoe and Vivian C. Deaton
    Report Number: NASA-TM-104269
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: High surface temperatures and temperature gradients can affect the vibratory characteristics and stability of aircraft structures. Aircraft designers are relying more on finite-element model analysis methods to ensure sufficient vehicle structural dynamic stability throughout the desired flight envelope. Analysis codes that predict these thermal effects must be correlated and verified with experimental data. This paper presents experimental modal data for aluminum, titanium, and fiberglass plates heated at uniform, nonuniform, and transient heating conditions. These data are compared with vibration analysis results for the same heating conditions. The data show the effect of heat on each plate's modal characteristics, a comparison of predicted and measured plate vibration frequencies, the measured modal damping, and the effect of modeling material property changes and thermal stresses on the accuracy of the analytical results at nonuniform and transient heating conditions.
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    Report Date: January 1993
    Keywords:      Hot structures; Modal analysis; Modal testing; Vibration analysis
    Notes: Also presented at the SEM Structural Testing Technology at High Temperatures II Conference, Ojai, CA, Nov. 8-10, 1993.


  32. SIMULTANEOUS MEASUREMENT OF TEMPERATURE AND STRAIN USING FOUR CONNECTING WIRES , Technical Memorandum
    Authors: Allen R. Parker, Jr.
    Report Number: NASA-TM-104271
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper describes a new signal-conditioning technique for measuring strain and temperature which uses fewer connecting wires than conventional techniques. Simultaneous measurement of temperature and strain has been achieved by using thermocouple wire to connect strain gages to signal conditioning. This signal conditioning uses a new method for demultiplexing sampled analog signals and the Anderson current loop circuit. Theory is presented along with data to confirm that strain gage resistance change is sensed without appreciable error because of thermoelectric effects. Furthermore, temperature is sensed without appreciable error because of voltage drops caused by strain gage excitation current flowing through the gage resistance.
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    Subject Category: 35
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    Report Date: November 1993
    No. Pages: 12
    Funding Organization: WU 505-70-63
    Keywords:      Bridge circuits; Instrumentation; Signal demultiplexer; Simultaneous measurement; Strain gage measurements; Thermocouple; Transducers
    Notes: Presented at SEM Fall Conference and Exhibit — on Structural Testing Technology at High Temperature-II, Ojai, California, November 8, 1993.


  33. IDENTIFICATION OF INTEGRATED AIRFRAME-PROPULSION EFFECTS ON AN F-15 AIRCRAFT FOR APPLICATION TO DRAG MINIMIZATION , Technical Memorandum
    Authors: Gerard S. Schkolnik
    Report Number: NASA-TM-4532
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The application of an adaptive real-time measurement-based performance optimization technique is being explored for a future flight research program. The key technical challenge of the approach is parameter identification, which uses a perturbation-search technique to identify changes in performance caused by forced oscillations of the controls. The controls on the NASA F-15 highly integrated digital electronic control (HIDEC) aircraft were perturbed using inlet cowl rotation steps at various subsonic and supersonic flight conditions to determine the effect on aircraft performance. The feasibility of the perturbation-search technique for identify ing integrated airframe-propulsion system performance effects was successfully shown through flight experiments and postflight data analysis. Aircraft response and control data were analyzed postflight to identify gradients and to determine the minimum drag point. Changes in longitudinal acceleration as small as 0.004 g were measured, and absolute resolution was estimated to be 0.002 g or approximately 50 lbf of drag. Two techniques for identifying performance gradients were compared: a least-squares estimation algorithm and a modified maximum likelihood estimator algorithm. A complementary filter algorithm was used with the least squares estimator.
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    Subject Category: 07
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    Report Date: November 1993
    No. Pages: 26
    Funding Organization: WU-533-02-39
    Keywords:      Adaptive control; Aircraft flight tests; Aircraft performance; Flight optimization; Optimal control
    Notes: Also presented as AIAA 93-3764 for the Guidance, Navigation, and Control Conference held August 9-11, 1993 in Monterey, California


  34. MECHANICAL AND THERMAL BUCKLING ANALYSIS OF SANDWICH PANELS UNDER DIFFERENT EDGE CONDITIONS
    Authors: William L. Ko
    Report Number: NASA-TM-4535
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: By using the Rayleigh-Ritz method of minimizing the total potential energy of a structural system, combined load (mechanical or thermal load) buckling equations are established for orthotropic rectangular sandwich panels supported under four different edge conditions. Two-dimensional buckling interaction curves and three-dimensional buckling interaction surfaces are constructed for high-temperature honeycomb-core sandwich panels supported under four different edge conditions. The interaction surfaces provide easy comparison of the panel buckling strengths and the domains of symmetrical and antisymmetrical buckling associated with the different edge conditions. Thermal buckling curves of the sandwich panels also are presented. The thermal buckling conditions for the cases with and without thermal moments were found to be identical for the small deformation theory. In sandwich panels, the effect of transverse shear is quite large, and by neglecting the transverse shear effect, the buckling loads could be overpredicted considerably. Clamping of the edges could greatly increase buckling strength more in compression than in shear.
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    Report Date: January 1993
    Keywords:      Buckling interaction curves; Buckling interaction surfaces; Mechanical buckling; Sandwich panels; Thermal buckling
    Notes: Also presented at PICAST-1 Conference on Aerospace Science and Technology, Taiwan, Republic of China, Dec. 6-9, 1993.


  35. PRELIMINARY ANALYSIS FOR A MACH 8 CROSSFLOW TRANSITION EXPERIMENT ON THE PEGASUS SPACE BOOSTER , Technical Memorandum
    Authors: Leslie Gong, W. Lance Richards, Richard C. Monaghan and Robert D. Quinn
    Report Number: NASA-TM-104272
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A boundary-layer transition experiment is proposed for a future flight mission of the air-launched Pegasus space booster. The flight experiment requires attaching a glove assembly to the wing of the first-stage booster. The glove design consists of a spring and hook attachment system which allows for thermal growth of a steel 4130 skin. This paper presents results from one- and two-dimensional thermal analyses of the initial design. These analyses were performed to ensure the integrity of the wing and to define optimal materials for use in the glove. Results obtained from the thermal analysis using turbulent flow conditions showed a maximum temperature of approximately 305 degrees C (581 degrees F) and a chordwise temperature gradient of less than 8.9 degrees C/cm (40.5 degrees F/in.) for the critical areas in the upper glove skin. The temperatures obtained from these thermal analyses are well within the required temperature limits of the glove.
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    Subject Category: 01
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    Report Date: November 1993
    No. Pages: 22
    Funding Organization: WU 505-70-91
    Keywords:      Heat transfer; Pegasus; Thermal analysis; Thermostructures; Transition
    Notes: Presented at the Society for Experimental Mechanics, Structural Testing Technology at High Technology at High Temperature-II Conference, Ojai, CA, Nov. 8-10, 1993.


  36. MULTIDISCIPLINARY AEROELASTIC ANALYSIS OF A GENERIC HYPERSONIC VEHICLE , Technical Memorandum
    Authors: K. K. Gupta and K. L. Petersen
    Report Number: NASA-TM-4544
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents details of a flutter and stability analysis of aerospace structures such as hypersonic vehicles. Both structural and aerodynamic domains are discretized by the common finite element technique. A vibration analysis is first performed by the STARS code employing a block Lanczos solution scheme. This is followed by the generation of a linear aerodynamic grid for subsequent linear flutter analysis within subsonic and supersonic regimes of the flight envelope; the doublet lattice and constant pressure techniques are employed to generate the unsteady aerodynamic forces. Flutter analysis is then performed for several representative flight points. The nonlinear flutter solution if effected by first implementing a CFD solution of the entire vehicle. Thus, a 3-D unstructured grid for the entire flow domain is generated by a moving front technique. A finite element Euler solution is then implemented employing a quasi-implicit as well as an explicit solution scheme. A novel multidisciplinary analysis is the next effect that employs modal and aerodynamic data to yield aerodynamic damping characteristics. Such analyses are performed for a number of flight points to yield a large set of pertinent data that define flight flutter characteristics of the vehicle. This paper outlines the finite-element-based integrated analysis procedures in detail, which is followed by the results of numerical analyses of flight flutter simulation.
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    Subject Category: 15
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    Report Date: October 1993
    No. Pages: 14
    Funding Organization: WU 505-70
    Keywords:      Computational fluid dynamics; flutter and stability analysis; Generic hypersonic vehicles; Multidisciplinary aeroelastic analysis
    Notes: This paper was originally prepared as AIAA-93-5028 for the AIAA 5th International Aerospace Planes Conference, November 30-December 3, 1993, Munich, Germany.


  37. STRAIN GAGE MEASUREMENT ERRORS IN THE TRANSIENT HEATING OF STRUCTURAL COMPONENTS , Technical Memorandum
    Authors: W. Lance Richards
    Report Number: NASA-TM-104274
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Significant strain-gage errors may exist in measurements acquired in transient thermal environments if conventional correction methods are applied. Conventional correction theory was modified and a new experimental method was developed to correct indicated strain data for errors created in radiant heating environments ranging from 0.6 degrees C/sec (1 degree F/sec) to over 56 degrees C/sec (100 degrees F/sec). In some cases the new and conventional methods differed by as much as 30 percent. Experimental and analytical results were compared to demonstrate the new technique. For heating conditions greater than 6 degrees C/sec (10 degrees F/sec), the indicated strain data corrected with the developed technique compared much better to analysis than the same data corrected with the conventional technique.
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    Subject Category: 39
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    Report Date: December 1993
    No. Pages: 16
    Funding Organization: WU 505-70-63
    Keywords:      Apparent strain; Measurement errors; Strain–gauge measurement; Structural testing; Thermal stress analysis
    Notes: Also presented at the SEM Fall Connference and Exhibit: Structural Testing at High Temperature II, Ojai, CA, Nov. 8-10, 1993.


  38. THERMAL-FLUID ANALYSIS OF THE FILL AND DRAIN OPERATIONS OF A CRYOGENIC FUEL TANK , Technical Memorandum
    Authors: Craig A. Stephens, Gregory J. Hanna and Leslie Gong
    Report Number: NASA-TM-104273
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The Generic Research Cryogenic Tank was designed to establish techniques for testing and analyzing the behavior of reusable fuel tank structures subjected to cryogenic fuels and aerodynamic heating. The Generic Research Cryogenic Tank tests will consist of filling a pressure vessel to a prescribed fill level, waiting for steady-state conditions, then draining the liquid while heating the external surface to simulate the thermal environment associated with hypersonic flight. Initial tests of the Generic Research Cryogenic Tank will use liquid nitrogen with future tests requiring liquid hydrogen. Two-dimensional finite-difference thermal-fluid models were developed for analyzing the behavior of the Generic Research Cryogenic Tank during fill and drain operations. The development and results of the two-dimensional fill and drain models, using liquid nitrogen, are provided, along with results and discussion on extrapolating the model results to the operation of the full-size Generic Research Cryogenic Tank. These numerical models provided a means to predict the behavior of the Generic Research Cryogenic Tank during testing and to define the requirements for the Generic Research Cryogenic Tank support systems such as vent, drain, pressurization, and instrumentation systems. In addition, the fill model provided insight into the unexpected role of circumferential conduction in cooling the Generic Research Cryogenic Tank pressure vessel during fill operations.
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    Subject Category: 34
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    Report Date: December 1993
    No. Pages: 21
    Funding Organization: 505-70-63
    Keywords:      Aerodynamic heating; Cryogenics; Fuel tanks; Heat transfer; Liquid nitrogen; Thermal analysis; Thermal insulation
    Notes: Also prepared as paper for SEM Structural Testing Technology at High Temperatures II Conf., Ojai, CA, Nov. 8-10, 1993.


  39. COMPUTATION OF HEAT TRANSFER FROM IMPINGING TURBULENT JETS , Conference Report
    Authors: B. H. Chang and A. F. Mills
    Report Number: H-2291
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An axisymmetric jet impinging on a flat plate was simulated using a low-Reynolds number k - e (epsilon) model. The accuracy of the numerical calculations was validated by comparing the predicted flowfields with experimental data for various nozzle-plate distance to jet diameter ratios (H/d) and Reynolds numbers. Heat transfer predictions for various H/d showed good agreement with experimental data in the wall jet region, and the complex behavior of heat transfer for small H/d is predicted. However, predictions in the neighborhood of the stagnation point were poor for small H/d. The results suggest that improved modelling in the equation for the dissipation rate of the turbulence kinetic energy is required.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
    Availability:

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    Report Date: May 1993
    No. Pages: 6
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: The 6th International Symposium on Transport Phenomena in Thermal Engineering, Seoul, Korea, 5/9/93, Vol 4, p.245-250. Work sponsored by NASA Dryden through UCLA, Flight Sys. Research Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  40. FLOW OF SUPERCRITICAL HYDROGEN IN A UNIFORMLY HEATED CIRCULAR TUBE , Conference Report
    Authors: B. Youn and A. F. Mills
    Report Number: H-2292
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Turbulent flow of supercritical hydrogen through a uniformly heated circular tube has been investigated using numerical methods, for the range of 4 x 10(5) (superscript 5) =/< Re =/< 3 x 10(6)(superscript 6), 5 =/< q(w) (wall heat flux) =/< 10 MW / m(2) (superscript 2), 30 =/< T(in) (temperature inlet) =/< 90 K, and 5 =/< P(in) (pressure inlet) =/< 15 MPa. The purpose is to validate a turbulence model and calculation method for the design of active cooling systems of hydrogen-fueled hypersonic aircraft, where the hydrogen fuel is used as coolant. The PHOENICS software package was used for the computations, which required special provision for evaluation of the thermophysical properties of the supercritical hydrogen, and a low Reynolds number form of the k - e turbulence model. Pressure drop and heat transfer data were compared with experiment and existing correlations, and good agreement was demonstrated. For the pressure range considered here a “thermal spike” was observed and shown to be due to the secondary peak in specific heat, rather than the primary peak.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
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    Report Date: June 1993
    No. Pages: 24
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Numerical Heat Transfer, Part A, Vol 24, (1993) pp. 1-24. Document is the result of work sponsored by NASA Dryden, performed through UCLA, Flight Systems Research Center. Contract monitor, R. Quinn; funding facilitated by K. Iliff.
 
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