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  1. AERODYNAMIC LIFT AND MOMENT CALCULATIONS USING A CLOSED-FORMSOLUTION OF THE POSSIO EQUATION , Technical Memorandum
    Authors: Jensen Lin and Kenneth W. Iliff
    Report Number: NASA-TM-2000-209019
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In this paper, we present closed-form formulas for the lift and moment coefficients of a lifting surface in two-dimensional, unsteady, compressible, subsonic flow utilizing a newly developed explicit analytical solution of the Possio equation. Numerical calculations are consistent with previous numerical tables based on series expansions or ad hoc numerical schemes. More importantly, these formulas lend themselves readily to flutter analysis, compared with the tedious table-look-up schemes currently in use.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (423 KBytes)
          PDF (311 KBytes)
    Report Date: April 2000
    No. Pages: 28
    Keywords:      Aeroelasticity; Lift and moment coefficients; Possio equation; Subsonicaerodynamics; Unsteady aerodynamics
    Notes: Jensen Lin, University of California, Los Angeles, California and Kenneth W. Iliff, Dryden Flight Research Center, Edwards, California. Part of NASA DFRC grant NCC2-374 with UCLA.


  2. USER'S GUIDE FOR COMPUTER PROGRAM THAT ROUTES SIGNAL TRACES , Technical Memorandum
    Authors: David R. Hedgley, Jr.
    Report Number: NASA-TM-2000-209036
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This disk contains both a FORTRAN computer program and the corresponding user's guide that facilitates both its incorporation into your system and its utility. The computer program represents an efficient algorithm that routes signal traces on layers of a printed circuit with both through-pins and surface mounts. The computer program included is an implementation of the ideas presented in the theoretical paper titled A Formal Algorithm for Routing Signal Traces on a Printed Circuit Board, NASA TP-3639 published in 1996. The computer program in the 'connects' file can be read with a FORTRAN compiler and readily integrated into software unique to each particular environment where it might be used.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 61
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          PDF (608 KBytes)
    Report Date: December 2000
    No. Pages: 67
    Funding Organization: 529-50-04-T2RR-000
    Keywords:      Artificial intelligence; Printed circuit board; Routing


  3. ERAST: SCIENTIFIC APPLICATIONS AND TECHNOLOGY COMMERCIALIZATION , Conference Proceedings
    Authors: Compiled by John D. Hunley and Yvonne Kellogg
    Report Number: NASA-CP-2000-209031
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This is a conference publication for an event designed to inform potential contractors and appropriate personnel in various scientific disciplines that the ERAST vehicles have reached a certain level of maturity and are available to perform a variety of missions ranging from data gathering to telecommunications. There are multiple applications of the technology and a great many potential commercial and governmental markets. As high altitude platforms, the ERAST vehicles can gather data at higher resolution than satellites and can do so continuously, whereas satellites pass over a particular area only once each orbit.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 04 & 43
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          PDF (7,427 KBytes)
    Report Date: September 2000
    No. Pages: 203
    Funding Organization: WU 282-00-00-76-GJ-TT-0-00R-000
    Keywords:      Communications; Geodesy; High altitude; Remote sensing; Research aircraft
    Notes: Proceedings of addresses, sessions, and workshops of the NASA ERAST Exclusive Preview sponsored by NASA Dryden Flight Research Center, Edwards, California, October 13, 1999.


  4. IN-FLIGHT WING PRESSURE DISTRIBUTIONS FOR THE NASA F/A-18A HIGH ALPHA RESEARCH VEHICLE , Technical Paper
    Authors: Mark C. Davis and John A. Saltzman
    Report Number: NASA-TM-2000-209018
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Pressure distributions on the wings of the F/A-18A High Alpha Research Vehicle (HARV) were obtained using both flush-mounted pressure orifices and surface- mounted pressure tubing. During quasi-stabilized 1-g flight, data were gathered at ranges for angle of attack from 5 degrees to 70 degrees, for angle of sideslip from - 12 degrees to +12 degrees, and for Mach from 0.23 to 0.64, at various engine settings, and with and without the leading edge extension fence installed. Angle of attack strongly influenced the wing pressure distribution, as demonstrated by a distinct flow separation pattern that occurred between the range from 15 degrees to 30 degrees. Influence by the leading edge extension fence was evident on the inboard wing pressure distribution, but little influence was seen on the outboard portion of the wing. Angle-of- sideslip influence on wing pressure distribution was strongest at low angle of attack. Influence of Mach number was observed in the regions of local supersonic flow, diminishing as angle of attack was increased. Engine throttle setting had little influence on the wing pressure distribution.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02, 05
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          Postscript (2,601 KBytes)
          PDF (1,239 KBytes)
    Report Date: March 2000
    No. Pages: 63
    Keywords:      F-18; Flight test; Flow visualization; Lift; Pressure distribution; Separation; Wingpressures


  5. DEVELOPMENT OF A FLUSH AIRDATA SENSING SYSTEM ON A SHARP-NOSED VEHICLE FOR FLIGHT AT MACH 3 TO 8 , Conference Paper
    Authors: Mark C. Davis, Joseph W. Pahle, John Terry White, Laurie A. Marshall, Michael J. Mashburn and Rick Franks
    Report Number: 2000-0504
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden Flight Research Center has developed a flush airdata sensing (FADS) system on a sharp-nosed, wedge-shaped vehicle. This paper details the design and calibration of a real-time angle-of-attack estimation scheme developed to meet the onboard airdata measurement requirements for a research vehicle equipped with a supersonic-combustion ramjet engine. The FADS system has been designed to perform in flights at Mach 3- 8 and at -6 deg - 12 deg angle of attack. The description of the FADS architecture includes port layout, pneumatic design, and hardware integration. Predictive models of static and dynamic performance are compared with wind-tunnel results across the Mach and angle-of-attack range. Results indicate that static angle-of-attack accuracy and pneumatic lag can be adequately characterized and incorporated into a real-time algorithm.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
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          Postscript (1,861 KBytes)
          PDF (712 KBytes)
    Report Date: January 2000
    No. Pages: 16
    Funding Organization: WU 522-51-54-00-50-00-X43
    Keywords:      Airdata calibration; FADS; Flush airdata sensing system; Hypersonics; Wedge forebody;
    Notes: Paper presented at 38th AIAA Aerospace Sciences Meeting and Exhibit, 10-13 January 2000, Reno, NV, AIAA 2000- 0504. M. Davis, J. Pahle, J. White and L. Marshall of NASA Dryden Flight Research Center, Edwards, CA. M. Mashburn of Micro Craft, Inc., Tullahoma, TN. Rick Franks of Sverdrup Corp., Arnold AFB, TN.


  6. DESIGN AND EVALUATION OF A NEW BOUNDARY-LAYER RAKE FOR FLIGHT TESTING , Technical Memorandum
    Authors: Trong T. Bui, David L. Oates and Jose C. Gonsalez
    Report Number: NASA-TM-2000-209014
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A new boundary-layer rake has been designed and built for flight testing on the NASA Dryden Flight Research Center F-15B/Flight Test Fixture. A feature unique to this rake is its curved body, which allows pitot tubes to be more densely clustered in the near-wall region than conventional rakes allow. This curved rake design has a complex three-dimensional shape that requires innovative solid-modeling and machining techniques. Finite-element stress analysis of the new design shows high factors of safety. The rake has passed a ground test in which random vibration measuring 12 g rms was applied for 20 min in each of the three normal directions. Aerodynamic evaluation of the rake has been conducted in the NASA Glenn Research Center 8 x 6 Supersonic Wind Tunnel at Mach 0-2. The pitot pressures from the new rake agree with conventional rake data over the range of Mach numbers tested. The boundary-layer profiles computed from the rake data have been shown to have the standard logarithmic-law profile. Skin friction values computed from the rake data using the Clauser plot method agree with the Preston tube results and the van Driest II compressible skin friction correlation to approximately ±5 percent.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02, 34
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          Postscript (509 KBytes)
          PDF (239 KBytes)
    Report Date: January 2000
    No. Pages: 21
    Funding Organization: WU 529-50-04-T2-RR-00-000
    Keywords:      Boundary layer; Compressible flows; Flight testing; Pitot pressure rake; Skin friction
    Notes: Presented at 38th AIAA Aerospace Sciences Conference, Reno, Nevada, January 10-13, 2000, AIAA-2000-0503. Trong T. Bui and David L. Oates of NASA Dryden Flight Research Center; Jose C. Gonsalez, Dynacs Engineering Co., Inc., Brookpark, Ohio.


  7. A REAL-TIME METHOD FOR ESTIMATING VISCOUS FOREBODY DRAG COEFFICIENTS , Technical Memorandum
    Authors: Stephen A. Whitmore, Marco Hurtado, Jose Rivera and Jonathan W. Naughton
    Report Number: NASA-TM-2000-209015
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper develops a real-time method based on the law of the wake for estimating forebody skin-friction coefficients. The incompressible law-of-the-wake equations are numerically integrated across the boundary layer depth to develop an engineering model that relates longitudinally averaged skin-friction coefficients to local boundary layer thickness. Solutions applicable to smooth surfaces with pressure gradients and rough surfaces with negligible pressure gradients are presented. Model accuracy is evaluated by comparing model predictions with previously measured flight data. This integral law procedure is beneficial in that skin-friction coefficients can be indirectly evaluated in real-time using a single boundary layer height measurement. In this concept a reference pitot probe is inserted into the flow, well above the anticipated maximum thickness of the local boundary layer. Another probe is servomechanism-driven and floats within the boundary layer. A controller regulates the position of the floating probe. The measured servomechanism position of this second probe provides an indirect measurement of both local and longitudinally averaged skin friction. Simulation results showing the performance of the control law for a noisy boundary layer are then presented.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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          Postscript (630 KBytes)
          PDF (360 KBytes)
    Report Date: January 2000
    No. Pages: 34
    Keywords:      Base drag; Boundary layer; Hypersonic vehicles; Skin friction
    Notes: Presented as AIAA 2000-0781 at the 38th Aerospace Sciences Meeting and Exhibit, January 10-13, 2000, Reno, Nevada. Stephen A. Whitmore and Marco Hurtado, NASA Dryden Flight Research Center, Edwards, CA; Jose Rivera, University of South Florida, Tampa, FL; and Jonathan W. Naughton, University of Wyoming, Laramie, WY.


  8. PEGASUS® WING-GLOVE EXPERIMENT TO DOCUMENT HYPERSONIC CROSSFLOW TRANSITION—MEASUREMENT SYSTEM AND SELECTED FLIGHT RESULTS , Technical Memorandum
    Authors: Arild Bertelrud, Geva de la Tova, Philip J. Hamory, Ronald Young, Gregory K. Noffz, Michael Dodson, Sharon S. Graves, John K. Diamond, James E. Bartlett, Robert Noack and David Knoblock
    Report Number: NASA-TM-2000-209016
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In a recent flight experiment to study hypersonic crossflow transition, boundary layer characteristics were documented. A smooth steel glove was mounted on the first stage delta wing of Orbital Sciences Corporation's Pegasus® launch vehicle and was flown at speeds of up to Mach 8 and altitudes of up to 250,000 ft. The wing-glove experiment was flown as a secondary payload off the coast of Florida in October 1998. This paper describes the measurement system developed. Samples of the results obtained for different parts of the trajectory are included to show the characteristics and quality of the data. Thermocouples and pressure sensors (including Preston tubes, Stanton tubes, and a ‘probeless' pressure rake showing boundary layer profiles) measured the time-averaged flow. Surface hot-films and high-frequency pressure transducers measured flow dynamics. Because the vehicle was not recoverable, it was necessary to design a system for real-time onboard processing and transmission. Onboard processing included spectral averaging. The quality and consistency of data obtained was good and met the experiment requirements.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
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          Postscript (1,300 KBytes)
          PDF (625 KBytes)
    Report Date: January 2000
    No. Pages: 30
    Keywords:      Boundary layer transition; Cross flow; Fault trees; Flight test instruments; Hypersonics; Onboard data processing; Pegasus air-launched booster
    Notes: Presented as AIAA-2000-0505 at the 38th Aerospace Sciences Meeting and Exhibit, January 10-13, 2000, Reno, NV. A. Bertelrud, AS&M, Inc., Edwards, CA; G. de la Tova, CSC, Edwards, CA; P. J. Hamory, R. Young, G. K. Noffz and M. Dodson, NASA Dryden Flight Research Center, Edwards, CA; S. S. Graves, J. K. Diamond, and J. E. Bartlett, NASA Langley Research Center, Hampton, VA; B. Noack, WYLE Laboratories, Hampton, VA; and D. Knoblock, NASA Kennedy Space Flight Center, Cape Canaveral, FL.


  9. DEVELOPMENT OF A FLUSH AIRDATA SENSING SYSTEM ON A SHARP-NOSED VEHICLE FOR FLIGHT AT MACH 3 TO 8 , Technical Memorandum
    Authors: Mark C. Davis , Joseph W. Pahle , John Terry White , Laurie A. Marshall , Michael J. Mashburn and Rick Franks
    Report Number: NASA-TM-2000-209017
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden Flight Research Center has developed a flush airdata sensing (FADS) system on a sharp-nosed, wedge-shaped vehicle. This paper details the design and calibration of a real-time angle-of-attack estimation scheme developed to meet the onboard airdata measurement requirements for a research vehicle equipped with a supersonic-combustion ramjet engine. The FADS system has been designed to perform in flights at speeds between Mach 3 and Mach 8 and at angles of attack between -6 deg and 12 deg. The description of the FADS architecture includes port layout, pneumatic design, and hardware integration. Predictive models of static and dynamic performance are compared with wind- tunnel results across the Mach and angle-of-attack range. Results indicate that static angle-of-attack accuracy and pneumatic lag can be adequately characterized and incorporated into a real-time algorithm.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
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          Postscript (2,112 KBytes)
          PDF (813 KBytes)
    Report Date: May 2000
    No. Pages: 26
    Funding Organization: WU 522-51-54-00-50-00-X43
    Notes: Author modification of paper presented at 38th AIAA Aerospace Sciences Meeting and Exhibit, 10-13 January 2000, Reno, NV, AIAA 2000-0504. M. Davis, J. Pahle, J. White and L. Marshall of NASA Dryden Flight Research Center, Edwards, CA; M. Mashburn of Micro Craft, Inc., Tullahoma, TN; and Rick Franks of Sverdrup Corp., Arnold AFB, TN.


  10. LONGITUDINAL HANDLING QUALITIES OF THE TU-144LL AIRPLANE AND COMPARISONS WITH OTHER LARGE, SUPERSONIC AIRCRAFT , Technical Memorandum
    Authors: Timothy H. Cox and Alisa Marshall
    Report Number: NASA-TM-2000-209020
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Four flights have been conducted using the Tu-144LL supersonic transport aircraft with the dedicated objective of collecting quantitative data and qualitative pilot comments. These data are compared with the following longitudinal flying qualities criteria: Neal-Smith, short-period damping, time delay, control anticipation parameter, phase delay, pitch bandwidth as a function of time delay, and flightpath as a function of pitch bandwidth. Determining the applicability of these criteria and gaining insight into the flying qualities of a large, supersonic aircraft are attempted. Where appropriate, YF-12, XB-70, and SR-71 pilot ratings are compared with the Tu-144LL results to aid in the interpretation of the Tu-144LL data and to gain insight into the application of criteria. The data show that approach and landing requirements appear to be applicable to the precision flightpath control required for up-and-away flight of large, supersonic aircraft. The Neal-Smith, control anticipation parameter, and pitch-bandwidth criteria tend to correlate with the pilot comments better than the phase delay criterion. The data indicate that the detrimental flying qualities implication of decoupled pitch-attitude and flightpath responses occurring for high-speed flight may be mitigated by requiring the pilot to close the loop on flightpath or vertical speed.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (860 KBytes)
          PDF (603 KBytes)
    Report Date: May 2000
    No. Pages: 45
    Funding Organization: WU 529-50-24-E8-RC-00-000
    Keywords:      Bandwidth criterion; Control anticipation criterion; Flying qualities; Neal-Smith criterion; Short period criteria; SR-71 aircraft; Supersonic flying qualities; TU-144 aircraft; XB-70 aircraft


  11. OPTIMAL PITCH THRUST-VECTOR ANGLE AND BENEFITS FOR ALL FLIGHT REGIMES , Technical Memorandum
    Authors: Glenn B. Gilyard and Alexander Bolonkin
    Report Number: NASA-TM-2000-209021
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA Dryden Flight Research Center is exploring the optimum thrust-vector angle on aircraft. Simple aerodynamic performance models for various phases of aircraft flight are developed and optimization equations and algorithms are presented in this report. Results of optimal angles of thrust vectors and associated benefits for various flight regimes of aircraft (takeoff, climb, cruise, descent, final approach, and landing) are given. Results for a typical wide-body transport aircraft are also given. The benefits accruable for this class of aircraft are small, but the technique can be applied to other conventionally configured aircraft. The lower L/D aerodynamic characteristics of fighters generally would produce larger benefits than those produced for transport aircraft.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02-01, 02-03, 03-01
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          Postscript (479 KBytes)
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    Report Date: March 2000
    No. Pages: 28
    Keywords:      Commercial aircraft; Efficiency; Lift-to-drag ratio; Optimization; Thrust vectoring
    Notes: Glenn B. Gilyard, NASA Dryden Flight Research Center, Edwards, California and Alexander Bolonkin, Senior Research Associate of the National Research Council, Edwards, California.


  12. EXPERIMENTAL AND NUMERICAL CHARACTERIZATION OF A STEADY-STATE CYLINDRICAL BLACKBODY CAVITY AT 1100 DEGREES CELSIUS , Technical Memorandum
    Authors: Thomas J. Horn and Amanie N. Abdelmessih
    Report Number: NASA-TM-2000-209022
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A blackbody calibration furnace at the NASA Dryden Flight Research Center is used to calibrate heat flux gages. These gages are for measuring the aerodynamic heat flux on hypersonic flight vehicle surfaces. The blackbody is a graphite tube with a midplane partition which divides the tube into two compartments (dual cavities). Electrical resistance heating is used to heat the graphite tube. This heating and the boundary conditions imposed on the graphite tube result in temperature gradients along the walls of the blackbody cavity. This paper describes measurements made during steady-state operation and development of finite- difference thermal models of the blockbody furnace at 1100 °C. Two configurations were studied, one with the blackbody outer surface insulated and the other without insulation. The dominant modes of heat transfer were identified for each configuration and the effect of variations in material properties and electric current that was passed through the blackbody were quantified.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
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          Postscript (483 KBytes)
          PDF (293 KBytes)
    Report Date: August 2000
    No. Pages: 21
    Funding Organization: WU 522-32-24-E8-RS-00-000
    Keywords:      Blackbody calibration furnace; Heat flux calibration; Heat transfer; Numerical
    Notes: Presented at 34th National Heat Transfer Conference, Pittsburgh, Pennsylvania, August 20-22, 2000, paper NHTC2000- 12140. Dr. Abdelmessih is an Associate Professor, Mechanical Engineering Department at St Martin's College, Lacey, Washington and a participant in the NASA/ASEE Summer Faculty Fellowship Program.


  13. THE SR-71 TEST BED AIRCRAFT:A FACILITY FOR HIGH-SPEED FLIGHT RESEARCH , Technical Publication
    Authors: Stephen Corda, Timothy R. Moes, Masashi Mizukami, Neal E. Hass, Daniel Jones, Richard C. Monaghan, Ronald J. Ray, Michele L. Jarvis and Nathan Palumbo
    Report Number: NASA-TP-2000-209023
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The SR-71 test bed aircraft is shown to be a unique platform to flight-test large experiments to supersonic Mach numbers. The test bed hardware mounted on the SR-71 upper fuselage is described. This test bed hardware is composed of a fairing structure called the 'canoe' and a large 'reflection plane' flat plate for mounting experiments. Total experiment weights, including the canoe and reflection plane, as heavy as 14,500 lb can be mounted on the aircraft and flight-tested to speeds as fast as Mach 3.2 and altitudes as high as 80,000 ft. A brief description of the SR-71 aircraft is given, including details of the structural modifications to the fuselage, modifications to the J58 engines to provide increased thrust, and the addition of a research instrumentation system. Information is presented based on flight data that describes the SR-71 test bed aerodynamics, stability and control, structural and thermal loads, the canoe internal environment, and reflection plane flow quality. Guidelines for designing SR-71 test bed experiments are also provided.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02, 05, 07, 08
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          Postscript (917 KBytes)
          PDF (781 KBytes)
    Report Date: June 2000
    No. Pages: 39
    Keywords:      Flight test; SR-71 aerodynamics; SR-71 aircraft; SR-71 stability and control; SR-71 upper fuselage flow field; Test bed aircraft.


  14. SUMMARY OF TRANSITION RESULTS FROM THE F-16XL-2 SUPERSONIC LAMINAR FLOW CONTROL EXPERIMENT , Meeting Paper
    Authors: Laurie A. Marshall
    Report Number: H-2410
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A variable-porosity suction glove has been flown on the F-16XL-2 aircraft to demonstrate the feasibility of this technology for the proposed High-Speed Civil Transport. Boundary-layer transition data on the titanium glove primarily have been obtained at speeds of Mach 2.0 and altitudes of 15,240-16,764 m (50,000-55,000 ft). The objectives of this flight experiment have been to achieve 0.50-0.60 chord laminar flow on a highly swept wing at supersonic speeds and to provide data to validate codes and suction design. The most successful laminar flow results have not been obtained at the glove design point, a speed of Mach 1.9 at an altitude of 15,240 m (50,000 ft); but rather at a speed of Mach 2.0 and an altitude of 16,154 m (53,000 ft). Laminar flow has been obtained to more than 0.46 wing chord at a Reynolds number of 22.7 X 10 to the sixth. A turbulence diverter has been used to initially obtain a laminar boundary layer at the attachment line. A lower-surface shock fence was required to block an inlet shock from the wing leading edge. This paper discusses research variables that directly impact the ability to obtain laminar flow and techniques to correct for these variables.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (1,240 KBytes)
          PDF (814 KBytes)
    Report Date: August 2000
    No. Pages: 13
    Keywords:      F-16XL; Hot-films; Laminar flow control; Suction; Transition.
    Notes: Presented at the AIAA August 2000 Conferences, Session APA-15 'Effectiveness of Flow Control Techniques,' Denver, Colorado, August 14-17, 2000, AIAA-2000-4418.


  15. USER'S GUIDE FOR SKETCH , Technical Memorandum
    Authors: David R. Hedgley, Jr.
    Report Number: NASA-TM-2000-210388
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A user's guide for the computer program SKETCH is presented on this disk. SKETCH solves a popular problem in computer graphics-the removal of hidden lines from images of solid objects. Examples and illustrations are included in the guide. Also included is the SKETCH program, so a user can incorporate the information into a particular software system.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 59
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          PDF (930 KBytes)
    Report Date: December 2000
    No. Pages: 72
    Funding Organization: 710550 4M 18E 8RR1 005
    Keywords:      Eulerian angles; Hidden line; ICORE; SKETCH


  16. 1999 RESEARCH ENGINEERING ANNUAL REPORT , Technical Memorandum
    Authors: Compiled by Edmund Hamlin, Everlyn Cruciani and Patricia Pearson
    Report Number: NASA-TM-2000-209026
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Selected research and technology activities at Dryden Flight Research Center are summarized. These activities exemplify the Center's varied and productive research efforts.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 99
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          Postscript (13,425 KBytes)
          PDF (4,318 KBytes)
    Report Date: August 2000
    No. Pages: 72
    Funding Organization: WU 529-50-04-MI-00-RR-0-005-000
    Keywords:      Aerodynamics; Flight; Flight controls; Flight systems; Flight test;Instrumentation; Propulsion; Structures; Structural dynamics


  17. FLOW FIELD SURVEY IN THE TEST REGION OF THE SR-71 AIRCRAFTTEST BED CONFIGURATION , Technical Memorandum
    Authors: Masashi Mizukami, Daniel Jones and Vladimir D. Weinstock
    Report Number: NASA-TM-2000-209025
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flat plate and faired pod have been mounted on a NASA SR-71A aircraft for use as a supersonic flight experiment test bed. A test article can be placed on the flat plate; the pod can contain supporting systems. A series of test flights has been conducted to validate this test bed configuration. Flight speeds to a maximum of Mach 3.0 have been attained. Steady-state sideslip maneuvers to a maximum of 2 degrees have been conducted, and the flow field in the test region has been surveyed. Two total-pressure rakes, each with two flow-angle probes, have been placed in the expected vicinity of an experiment. Static-pressure measurements have been made on the flat plate. At subsonic and low supersonic speeds with no sideslip, the flow in the surveyed region is quite uniform. During sideslip maneuvers, localized flow distortions impinge on the test region. Aircraft sideslip does not produce a uniform sidewash over the test region. At speeds faster than Mach 1.5, variable-pressure distortions were observed in the test region. Boundary-layer thickness on the flat plate at the rake was less than 2.1 in. For future experiments, a more focused and detailed flow-field survey than this one would be desirable.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          PDF (833 KBytes)
    Report Date: August 2000
    No. Pages: 122
    Keywords:      Flight testing; Flow distortion; Fluid flow; SR-71 aircraft; Supersonic inlet.
    Notes: Masashi Mizukami and Daniel Jones, NASA Dryden Flight Research Center, Edwards, California; and Vladimir D. Weinstock, Analytic Services and Materials, Inc., Hampton, Virginia. Equations (9)-(11) and (13)-(15) are included in a patent filed on a NASA invention.


  18. COMPARISON OF THREE WIND MEASURING SYSTEMS FOR FLIGHT TEST , Meeting Paper
    Authors: Edward H. Teets, Jr. and Philip O. Harvey
    Report Number: H-2415
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A preliminary field test of the accuracy of wind velocity measurements obtained using global positioning system-tracked rawinsonde balloons has been performed. Wind comparisons have been conducted using global positioning system (GPS) and radio automatic theodolite sounder (RATS) rawinsondes and a high-precision range instrumention radar-tracked reflector. Wind velocity differences between the GPS rawinsondes and the radar were significantly less than between the RATS rawinsondes and the radar. These limited test results indicate a root-mean-square wind velocity difference from 4.98 kn (2.56 m/sec) for the radar and RATS to 1.09 kn (0.56 m/sec) for the radar and GPS. Differences are influenced by user reporting requirements, data processing techniques, and the inherent tracking accuracies of the system. This brief field test indicates that the GPS sounding system tracking data are more precise than the RATS system. When high-resolution wind data are needed, use of GPS rawinsonde systems can reduce the burden on range radar operations.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 47, 46, 70, 02
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    Report Date: September 2000
    No. Pages: 6
    Keywords:      Atmospheric science; Flight test; GPS rawinsonde; RADAR; Rawinsonde.
    Notes: Presented at the American Meteorological Society's Ninth Conference on Aviation, Range, and Aerospace Meteorology, Orlando, Florida, September 11-15, 2000. Edward H. Teets, Jr., NASA Dryden Flight Research Center, Edwards, CA and Philip O. Harvey, Air Force Flight Test Center, Edwards Air Force Base, CA.


  19. FLYING QUALITY ANALYSIS OF A JAS 39 GRIPEN MINISTICK CONTROLLER IN AN F/A-18 AIRCRAFT , Technical Memorandum
    Authors: John F. Carter and P. C. Stoliker
    Report Number: NASA-TM-2000-209024
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden conducted a handling qualities experiment using a small displacement centerstick controller that Saab-Scania developed for the JAS 39 Gripen aircraft. The centerstick, or ministick, was mounted in the rear cockpit of an F/A-18 aircraft. Production support flight control computers (PSFCC) provided a pilot-selectable research control system. The objectives for this experiment included determining whether the mechanical characteristics of the centerstick controller had any significant effect on the handling qualities of the F/A-18, and determining the usefulness of the PSFCCs for this kind of experiment. Five pilots evaluated closed-loop tracking tasks, including echelon and column formation flight and target following. Cooper-Harper ratings and pilot comments were collected for each maneuver. This paper describes the test system, including the PSFCCs, the Gripen centerstick, and the flight test experiment. The paper presents results of longitudinal handling qualities maneuvers, including low order equivalent systems, Neal-Smith, and controls anticipation parameter analyses. The experiment showed that, while the centerstick controller provided a different aircraft feel, few handling qualities deficiencies resulted. It also demonstrated that the PSFCCs were useful for this kind of investigation.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
    Availability:
        Format(s) on-line:
          Postscript (500 KBytes)
          PDF (359 KBytes)
    Report Date: August 2000
    No. Pages: 21
    Funding Organization: WU 529 61 14 M1 00 14 0 00 S 000
    Keywords:      Aircraft control; Flight control; Handling qualities; JAS 39 Gripen; Piloting
    Notes: Presented at AIAA Guidance Navigation and Control Conference, Denver, Colorado, August 14-17, 2000, AIAA-2000-4444.


  20. THERMOELASTIC ANALYSIS OF HYPER-X CAMERA WINDOWS SUDDENLY EXPOSED TO MACH 7 STAGNATION AEROTHERMAL SHOCK , Technical Publication
    Authors: William L. Ko and Leslie Gong
    Report Number: NASA-TP-2000-209030
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: To visually record the initial free flight event of the Hyper-X research flight vehicle immediately after separation from the Pegasus® booster rocket, a video camera was mounted on the bulkhead of the adapter through which Hyper-X rides on Pegasus. The video camera was shielded by a protecting camera window made of heat-resistant quartz material. When Hyper-X separates from Pegasus, this camera window will be suddenly exposed to Mach 7 stagnation thermal shock and dynamic pressure loading (aerothermal loading). To examine the structural integrity, thermoelastic analysis was performed, and the stress distributions in the camera windows were calculated. The critical stress point where the tensile stress reaches a maximum value for each camera window was identified, and the maximum tensile stress level at that critical point was found to be considerably lower than the tensile failure stress of the camera window material.
    Distribution/Availability: Unclassified - Unlimited
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          PDF (451 KBytes)
    Report Date: September 2000
    No. Pages: 29
    Keywords:      Circular disk; Dynamic pressure bending; Mach 7 heating; Mechanical stresses;


  21. VISUALIZATION OF IN-FLIGHT FLOW PHENOMENA USING INFRARED THERMOGRAPHY , Technical Memorandum
    Authors: D. W. Banks, C. P. van Dam, H. J. Shiu and G. M. Miller
    Report Number: NASA-TM-2000-209027
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Infrared thermography was used to obtain data on the state of the boundary layer of a natural laminar flow airfoil in supersonic flight. In addition to the laminar-to-turbulent transition boundary, the infrared camera was able to detect shock waves and present a time dependent view of the flow field. A time dependent heat transfer code was developed to predict temperature distributions on the test subject and any necessary surface treatment. A commercially available infrared camera was adapted for airborne use in this application. Readily available infrared technology has the capability to provide detailed visualization of various flow phenomena in subsonic to hypersonic flight regimes.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
    Availability:
        Format(s) on-line:
          Postscript (941 KBytes)
          PDF (385 KBytes)
    Report Date: July 2000
    No. Pages: 17
    Funding Organization: WU 529-35-14-M1-00-38-0-00-S-000
    Keywords:      Boundary layer transition; Flow visualization; Infrared thermography; Shockwaves; Supersonic flow
    Notes: Paper no. 24, prepared for the International Symposium on Flow Visualization, August 22-25, 2000, Edinburgh, Scotland. D. W. Banks, NASA Dryden Flight Research Center, Edwards, CA; C. P. van Dam and H. J. Shiu, Dept. of Mechanical and Aeronautical Eng., Univ. of California, Davis, CA; and G. M. Miller, PVP Advanced EP Systems, Orange, CA.


  22. AN EVALUATION TECHNIQUE FOR AN F/A-18 AIRCRAFT LOADS MODEL USING F/A-18 SYSTEMS RESEARCH AIRCRAFT FLIGHT DATA , Technical Memorandum
    Authors: Candida D. Olney, Heather Hillebrandt and Eric Y. Reichenbach
    Report Number: NASA-TM-2000-209028
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A limited evaluation of the F/A-18 baseline loads model was performed on the Systems Research Aircraft at NASA Dryden Flight Research Center (Edwards, California). Boeing developed the F/A-18 loads model using a linear aeroelastic analysis in conjunction with a flight simulator to determine loads at discrete locations on the aircraft. This experiment was designed so that analysis of doublets could be used to establish aircraft aerodynamic and loads response at 20 flight conditions. Instrumentation on the right outboard leading edge flap, left aileron, and left stabilator measured the hinge moment so that comparisons could be made between in-flight-measured hinge moments and loads model-predicted values at these locations. Comparisons showed that the difference between the loads model-predicted and in-flight-measured hinge moments was up to 130 percent of the flight limit load. A stepwise regression technique was used to determine new loads derivatives. These derivatives were placed in the loads model, which reduced the error to within 10 percent of the flight limit load. This paper discusses the flight test methodology, a process for determining loads coefficients, and the direct comparisons of predicted and measured hinge moments and loads coefficients.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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          Postscript (590 KBytes)
          PDF (372 KBytes)
    Report Date: July 2000
    No. Pages: 30
    Funding Organization: WU 529-61-14-E8-14-00-AAW
    Keywords:      Active aeroelastic wing; F/A-18; Flight test; Loads; Simulation30
    Notes: Prepared for the 31st SFTE Annual Symposium 2000, Torino, Italy, September 18-22, 2000. Candida D. Olney and Heather Hillebrandt, NASA Dryden Flight Research Center, Edwards, California, and Eric Y. Reichenbach, The Boeing Company, Phantom Works, St. Louis, Missouri.


  23. RESULTS FROM F-18B STABILITY AND CONTROL PARAMETER ESTIMATION FLIGHT TESTS AT HIGH DYNAMIC PRESSURES.
    Authors: Timothy R. Moes, Gregory K. Noffz and Kenneth W. Iliff
    Report Number: NASA-TP-2000-209033
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A maximum-likelihood output-error parameter estimation technique has been used to obtain stability and control derivatives for the NASA F-18B Systems Research Aircraft. This work has been performed to support flight testing of the active aeroelastic wing (AAW) F-18A project. The goal of this research is to obtain baseline F-18 stability and control derivatives that will form the foundation of the aerodynamic model for the AAW aircraft configuration. Flight data have been obtained at Mach numbers between 0.85 and 1.30 and at dynamic pressures ranging between 600 and 1500 lbf/ft2. At each test condition, longitudinal and lateral-directional doublets have been performed using an automated onboard excitation system. The doublet maneuver consists of a series of single-surface inputs so that individual control-surface motions cannot be correlated with other control-surface motions. Flight test results have shown that several stability and control derivatives are significantly different than prescribed by the F-18B aerodynamic model. This report defines the parameter estimation technique used, presents stability and control derivative results, compares the results with predictions based on the current F-18B aerodynamic model, and shows improvements to the nonlinear simulation using updated derivatives from this research.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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          PDF (3,953 KBytes)
    Report Date: November 2000
    No. Pages: 143
    Keywords:      Active aeroelastic wing; Control derivatives; Maximum likelihood estimates; Parameter identification; Stability derivatives


  24. FLIGHT TEST EXPERIENCE WITH AN ELECTROMECHANICAL ACTUATOR ON THE F-18 SYSTEMS RESEARCH AIRCRAFT , Conference Paper
    Authors: Stephen C. Jensen, Gavin D. Jenney, Bruce Raymond and David Dawson
    Report Number: H-2425
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Development of reliable power-by-wire actuation systems for both aeronautical and space applications has been sought recently to eliminate hydraulic systems from aircraft and spacecraft and thus improve safety, efficiency, reliability, and maintainability. The Electrically Powered Actuation Design (EPAD) program was a joint effort between the Air Force, Navy, and NASA to develop and fly a series of actuators validating power-by-wire actuation technology on a primary flight control surface of a tactical aircraft. To achieve this goal, each of the EPAD actuators was installed in place of the standard hydraulic actuator on the left aileron of the NASA F/A-18B Systems Research Aircraft (SRA) and flown throughout the SRA flight envelope. Numerous parameters were recorded, and overall actuator performance was compared with the performance of the standard hydraulic actuator on the opposite wing. This paper discusses the integration and testing of the EPAD electromechanical actuator (EMA) on the SRA. The architecture of the EMA system is discussed, as well as its integration with the F/A-18 Flight Control System. The flight test program is described, and actuator performance is shown to be very close to that of the standard hydraulic actuator it replaced. Lessons learned during this program are presented and discussed, as well as suggestions for future research.
    Distribution/Availability: Unclassified - Unlimited
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          PDF (558 KBytes)
    Report Date: October 2000
    No. Pages: 11
    Funding Organization: 529-61-14-E8-14-00-SRA
    Notes: Presented at the 19th Digital Avionics Systems Conference, October 7-13, 2000, Philadelphia, Pennsylvania. G. Jenney and B. Raymond, Dynamic Controls, Inc., Dayton, Ohio; D. Dawson, Wright Laboratory, WPAFB, Ohio.


  25. A METHOD FOR CALCULATING TRANSIENT SURFACE TEMPERATURES AND SURFACE HEATING RATES FOR HIGH-SPEED AIRCRAFT , Technical Publication
    Authors: Robert D. Quinn and Leslie Gong
    Report Number: NASA-TP-2000-209034
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This report describes a method that can calculate transient aerodynamic heating and transient surface temperatures at supersonic and hypersonic speeds. This method can rapidly calculate temperature and heating rate time-histories for complete flight trajectories. Semi-empirical theories are used to calculate laminar and turbulent heat transfer coefficients and a procedure for estimating boundary-layer transition is included. Results from this method are compared with flight data from the X-15 research vehicle, YF-12 airplane, and the Space Shuttle Orbiter. These comparisons show that the calculated values are in good agreement with the measured flight data.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
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          PDF (343 KBytes)
    Report Date: December 2000
    No. Pages: 36
    Funding Organization: 529-35-34-E8-RR-00-000
    Keywords:      Aerodynamic heating; Boundary layer; Flight measurements; Flow hypersonic; Flow inviscid


  26. AIRBORNE COHERENT LIDAR FOR ADVANCED IN-FLIGHT MEASUREMENTS (ACLAIM) FLIGHT TESTING OF THE LIDAR SENSOR , Meeting Paper
    Authors: David C. Soreide, Rodney K. Bogue, L. J. Ehernberger, Stephen M. Hannon and David A. Bowdle
    Report Number: H-2428
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The purpose of the ACLAIM program is ultimately to establish the viability of light detection and ranging (lidar) as a forward-looking sensor for turbulence. The goals of this flight test are to: 1) demonstrate that the ACLAIM lidar system operates reliably in a flight test environment, 2) measure the performance of the lidar as a function of the aerosol backscatter coefficient (beta), 3) use the lidar system to measure atmospheric turbulence and compare these measurements to onboard gust measurements, and 4) make measurements of the aerosol backscatter coefficient, its probability distribution and spatial distribution. The scope of this paper is to briefly describe the ACLAIM system and present examples of ACLAIM operation in flight, including comparisons with independent measurements of wind gusts, gust-induced normal acceleration, and the derived eddy dissipation rate.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 03, 06
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          PDF (309 KBytes)
    Report Date: September 2000
    No. Pages: 8
    Keywords:      Atmospheric turbulence; Coherent lidar; Doppler lidar; Flight test; Optical radar; Structure function; Turbulence detection.
    Notes: Prepared for the American Meteorological Society 9th Conf. on Aviation, Range, and Aerospace Meteorology, Orlando, FL, Sept. 11-15, 2000. David Soreide, Boeing Co.; Rodney Bogue and L. J. Ehernberger, NASA Dryden Flight Research Center; Stephen Hannon, Coherent Technologies, Inc.; and David Bowdle, Univ. of Alabama at Huntsville.
 
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