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- F-16XL WING PRESSURE DISTRIBUTIONS AND SHOCK FENCE RESULTS FROM MACH 1.4 TO MACH 2.0 , Technical Memorandum
Authors: Stephen F. Landers, John A. Saltzman and Lisa J. Bjarke
Report Number: NASA-TM-97-206219
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Chordwise pressure distributions were obtained in-flight on the upper and lower surfaces of the F-16XL ship 2 aircraft wing between Mach 1.4 and Mach 2.0. This experiment was conducted to determine the location of shock waves which could compromise or invalidate a follow-on test of a large chord laminar flow control suction panel. On the upper surface, the canopy closure shock crossed an area which would be covered by a proposed laminar flow suction panel. At the laminar flow experiment design Mach number of 1.9, 91 percent of the suction panel area would be forward of the shock. At Mach 1.4, that value reduces to 65 percent. On the lower surface, a shock from the inlet diverter would impinge on the proposed suction panel leading edge. A chordwise plate mounted vertically to deflect shock waves, called a shock fence, was installed between the inlet diverter and the leading edge. This plate was effective in reducing the pressure gradients caused by
the inlet shock system.
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Subject Category: 02
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Report Date: October 1997
No. Pages: 43
Funding Organization: 529 31 24 00 24 00
Keywords: F-16XL aircraft; Supersonic transports; Shockwaves; Pressure distribution; Laminar flow
Notes: Stephen F. Landers and John A. Saltzman, PRC Inc., Edwards, California.Lisa J. Bjarke, NASA Dryden Flight Research Center, Edwards, California.
- FLIGHT WING SURFACE PRESSURE AND BOUNDARY-LAYER DATA REPORT FROM THE F-111 SMOOTH VARIABLE-CAMBER SUPERCRITICAL MISSION ADAPTIVE WING , Technical Memorandum
Authors: Sheryll Goecke Powers and Lannie D. Webb
Report Number: NASA-TM-4789
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg
or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4e + 06 to 6.5e + 06. Leading-edge cambers ranged from 0 deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.
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Subject Category: 02
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Report Date: June 1997
No. Pages: 861
Funding Organization: WU 953-36-00
Keywords: Boundary layers; Pressure distributions; Smooth variable camber; Supercritical wing
Notes: n.a.
- THERMOSTRUCTURAL BEHAVIOR OF A HYPERSONIC AIRCRAFT SANDWICH PANEL SUBJECTED TO HEATING ON ONE SIDE , Technical Memorandum
Authors: William L. Ko
Report Number: NASA-TM-4769
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Thermostructural analysis was performed on a heated titanium honeycomb-core sandwich panel. The sandwich panel was supported at its four edges with spar-like substructures that acted as heat sinks, which are generally not considered in the classical analysis. One side of the panel was heated to high temperature to simulate aerodynamic heating during hypersonic flight. Two types of surface heating were considered: (1) flat-temperature profile, which ignores the effect of edge heat sinks, and (2) dome-shaped-temperature profile, which approximates the actual surface temperature distribution associated with the existence of edge heat sinks. The finite-element method was used to calculate the deformation field and thermal stress distributions in the face sheets and core of the sandwich panel. The detailed thermal stress distributions in the sandwich panel are presented, and critical stress regions are identified. The study shows how the magnitudes of
those critical stresses and their locations change with different heating and edge conditions. This technical report presents comprehensive, three-dimensional graphical displays of thermal stress distributions in every part of a titanium honeycomb-core sandwich panel subjected to hypersonic heating on one side. The plots offer quick visualization of the structural response of the panel and are very useful for hot structures designers to identify the critical stress regions.
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Subject Category: 39
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Report Date: April 1997
No. Pages: 51
Funding Organization: WU 505-63-50-00-RS-00-000-01
Keywords: Heat-sink effect; One-sided heating; Sandwich panels; Thermal deformations; Thermal stress analysis; Thermal stress distributions
- COUPLING DYNAMICS IN AIRCRAFT: A HISTORICAL PERSPECTIVE , Special Publication
Authors: Richard E. Day
Report Number: NASA-SP-532
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Coupling dynamics can produce either adverse or beneficial stability and controllability, depending on the characteristics of the aircraft. This report presents archival anecdotes and analyses of coupling problems experienced by the X-series, Century series, and Space Shuttle aircraft. The three catastrophic sequential coupling modes of the X-2 airplane and the two simultaneous unstable modes of the X-15 and Space Shuttle aircraft are discussed. In addition, the most complex of the coupling interactions, inertia roll coupling, is discussed for the X-2, X-3, F-100A, and YF-102 aircraft. The mechanics of gyroscopics, centrifugal effect, and resonance in coupling dynamics are described. The coupling modes discussed are interacting multiple degrees of freedom of inertial and aerodynamic forces and moments. The aircraft are assumed to be rigid bodies. Structural couplings are not addressed. Various solutions for coupling instabilities are discussed.
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Report Date: March 1997
No. Pages: 70
Funding Organization: NASA SP-532
Keywords: Coupling dynamics; Coupling modes; F-100 aircraft; Space Shuttle; X-2 aircraft; X-3 aircraft; X-15 aircraft; YF-102 aircraft.
Notes: Richard E. Day (NASA retired). Prepared under NASA contract NASA 2-13445, subcontract no. TSD-93-RED-2805.
- DEVELOPMENT OF A CLOSED-LOOP STRAP DOWN ATTITUDE SYSTEM FOR AN ULTRAHIGH ALTITUDE FLIGHT EXPERIMENT , Technical Memorandum
Authors: Stephen A. Whitmore, Mike Fife and Logan Brashear
Report Number: NASA-TM-4775
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A low-cost attitude system has been developed for an ultrahigh altitude flight experiment. The experiment uses a remotely piloted sailplane, with the wings modified for flight at altitudes greater than 100,000 ft. Mission requirements deem it necessary to measure the aircraft pitch and bank angles with accuracy better than 1.0 deg and heading with accuracy better than 5.0 deg. Vehicle cost restrictions and gross weight limits make installing a commercial inertial navigation system unfeasible. Instead, a low-cost attitude system was developed using strap down components. Monte Carlo analyses verified that two vector measurements, magnetic field and velocity, are required to completely stabilize the error equations. In the estimating algorithm, body-axis observations of the airspeed vector and the magnetic field are compared against the inertial velocity vector and a magnetic-field reference model. Residuals are fed back to stabilize integration of
rate gyros. The effectiveness of the estimating algorithm was demonstrated using data from the NASA Dryden Flight Research Center Systems Research Aircraft (SRA) flight tests. The algorithm was applied with good results to a maximum 10 deg pitch and bank angles. Effects of wind shears were evaluated and, for most cases, can be safely ignored.
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Subject Category: 06
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Report Date: January 1997
No. Pages: 31
Funding Organization: WU 537-10-40
Keywords: Attitude estimation; Euler angles; Kalman filter; Quaternion; Ultrahigh altitude
Notes: Presented at the 35th Aerospace Sciences and Exhibit, January 6-9, 1997, Reno, Nevada.
- EXTRACTION OF LATERAL-DIRECTIONAL STABILITY AND CONTROL DERIVATIVES FOR THE BASIC F-18 AIRCRAFT AT HIGH ANGLES OF ATTACK , Technical Memorandum
Authors: Kenneth W. Iliff and Kon-Sheng Charles Wang
Report Number: NASA-TM-4786
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The results of parameter identification to determine the lateral-directional stability and control derivatives of an F-18 research aircraft in its basic hardware and software configuration are presented. The derivatives are estimated from dynamic flight data using a specialized identification program developed at NASA Dryden Flight Research Center. The formulation uses the linearized aircraft equations of motions in their continuous/discrete form and a maximum likelihood estimator that accounts for both state and measurement noise. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics, such as separated and vortical flows, over the aircraft. The derivatives are plotted as functions of angle of attack between 3 degrees and 47 degrees and compared with wind-tunnel predictions. The quality of the derivative estimates obtained by parameter identification is somewhat degraded because the maneuvers were flown with
the aircraft's control augmentation system engaged, which introduced relatively high correlations between the control variables and response variables as a result of control motions from the feedback control system.
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Report Date: February 1997
No. Pages: 39
Funding Organization: WU 505-68-50
Keywords: Aerodynamics; Flight-to-wind-tunnel comparisons; F-18 aircraft; High angle of attack; Maximum likelihood estimation; Parameter identification; Stability and control derivatives
Notes: None
- F/A-18A INLET FLOW CHARACTERISTICS DURING MANEUVERS WITH RAPIDLY CHANGING ANGLE OF ATTACK , Technical Memorandum
Authors: Andrew J. Yuhas , William G. Steenken , John G. Williams and Kevin R. Walsh
Report Number: NASA-TM-104327
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The performance and distortion levels of the right inlet of the F/A-18A High Alpha Research Vehicle were assessed during maneuvers with rapidly changing angle- of-attack at the NASA Dryden Flight Research Center, Edwards, California. The distortion levels were compared with those produced by current inlet-engine compatibility evaluation techniques. The objective of these analyses was to determine whether the results obtained for steady aerodynamic conditions were adequate to describe the inlet-generated distortion levels that occur during rapid aircraft maneuvers. The test data were obtained during 46 dynamic maneuvers at Mach numbers of 0.3 and 0.4. Levels of inlet recovery, peak dynamic circumferential distortion, and peak dynamic radial distortion of dynamic maneuvers for a General Electric F404-GE-400 turbofan engine were compared with estimations based on steady aerodynamic conditions. The comparisons were performed at equivalent
angle-of-attack, angle-of-sideslip, and Mach number. Results showed no evidence of peak inlet distortion levels being elevated by dynamic maneuver conditions at high angle-of-attack compared with steady aerodynamic estimations. During sweeps into high angle-of-attack, the peak distortion levels of the dynamic maneuvers rarely rose to steady aerodynamic estimations. The dynamic maneuvers were shown to be effective at identifying conditions when discrete changes in inlet behavior occur.
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Subject Category: 07
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Report Date: June 1997
No. Pages: 21
Funding Organization: WU 505-68-30
Keywords: Comparison to steady aerodynamic conditions; F/A-18A aircraft; High angle of
Notes: Oral presentation presented at the NASA High Angle-of-Attack Technology Conference, Inlet Aerodynamics and Propulsion Session, NASA Langley Research Center, September 17-19, 1996, Hampton, Virginia.
- A LABORATORY STUDY ON THE PHASE TRANSITION FOR POLAR STRATOSPHERIC CLOUD PARTICLES , Contractor Report
Authors: Edward H. Teets, Jr.
Report Number: NASA-CR-198056
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The nucleation and growth of different phases of simulated polar stratospheric cloud (PSC) particles were investigated in the laboratory. Solutions and mixtures of solutions at concentrations 1 to 5 m (molality) of ammonium sulfate, ammonium bisulfate, sodium chloride, sulfuric acid, and nitric acid were supercooled to prescribed temperatures below their equilibrium melting point. These solutions were contained in small diameter glass tubing of volumes ranging from 2.6 to 0.04 ml. Samples were nucleated by insertion of an ice crystal, or in some cases by a liquid nitrogen cooled wire. Crystallization velocities were determined by timing the crystal growth front passages along the glass tubing. Solution mixtures containing aircraft exhaust (soot) were also examined. Crystallization rates increased as "Delta"T(squared), where "Delta"T is the supercooling for weak solutions (2 m or less). The higher concentrated solutions (>3 m) showed rates
significantly less than "Delta"T(squared). This reduced rate suggested an onset of a glass phase. Results were applied to the nucleation of highly concentrated solutions at various stages of polar stratospheric cloud development within the polar stratosphere.
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Subject Category: 46
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Report Date: March 1997
No. Pages: 57
Funding Organization: WU 529-10-04
Keywords: Cloud research; Crystals; Stratospheric circulation; Temperature effects; Thermodynamic properties of liquids
Notes: NASA Contract NAG-2572
- THE F/A-18 HIGH-ANGLE-OF-ATTACK GROUND-TO-FLIGHT CORRELATION: LESSONS LEARNED , Technical Memorandum
Authors: Daniel W. Banks, David F. Fisher, Robert M. Hall, Gary E. Erickson, Daniel G. Murri, Sue B. Grafton and William G. Sewall
Report Number: NASA-TM-4783
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Detailed wind tunnel and flight investigations were performed on the F/A-18 configuration to explore the causes of many high-angle-of-attack phenomena and resulting disparities between wind tunnel and flight results at these conditions. Obtaining accurate predictions of full-scale flight aerodynamics from wind-tunnel tests is important and becomes a challenge at high-angle-of-attack conditions where large areas of vortical flow interact. The F/A-18 airplane was one of the first high-performance aircraft to have an unrestricted angle-of-attack envelope, and as such the configuration displayed many unanticipated characteristics. Results indicate that fixing forebody crossflow transition on models can result in a more accurate match of flow fields, and thus a more accurate prediction of aerodynamic characteristics of flight at high angles of attack. The wind tunnel results show that small geometry differences, specifically nosebooms and aft-end
distortion, can have a pronounced effect at high angles of attack and must be modeled in sub-scale tests in order to obtain accurate correlations with flight.
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Subject Category: 02
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Report Date: January 1997
No. Pages: 42
Funding Organization: WU 505-68-30
Keywords: Fighter aircraft; Forebody flows; High angle of attack; Vortex flows; Vortex interactions
Notes: Presented at NASA Langley High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia, September 17-19, 1996. Daniel Banks and David Fisher, Dryden Flight Research Center, Edwards, CA; Robert Hall, Gary Erickson, Daniel Murri, S
- STARS-AN INTEGRATED, MULTIDISCIPLINARY, FINITE-ELEMENT, STRUCTURAL, FLUIDS, AEROELASTIC, AND AEROSERVOELASTIC ANALYSIS COMPUTER PROGRAM , Technical Memorandum
Authors: K. K. Gupta
Report Number: NASA-TM-4795
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A multidisciplinary, finite element-based, highly graphics-oriented, linear and nonlinear analysis capability that includes such disciplines as structures, heat transfer, linear aerodynamics, computational fluid dynamics, and controls engineering has been achieved by integrating several new modules in the original STARS (STructural Analysis RoutineS) computer program. Each individual analysis module is general-purpose in nature and is effectively integrated to yield aeroelastic and aeroservoelastic solution of complex engineering problems. Examples of advanced NASA Dryden Flight Research Center projects analyzed by the code in recent years include the X-29A, F-18 High Alpha Research Vehicle/Thrust Vectoring Control System, B-52/Pegasus ®, Generic Hypersonics, National AeroSpace Plane (NASP), SR-71/Hypersonic Launch Vehicle, and High Speed Civil Transport (HSCT) projects. Extensive graphics capabilities exist for convenient model development and
postprocessing of analysis results. The program is written in modular form in standard FORTRAN language to run on a variety of computers, such as the IBM RISC/6000, SGI, DEC, Cray, and personal computer; associated graphics codes use OpenGL and IBM/graPHIGS language for color depiction. This program is available from COSMIC, the NASA agency for distribution of computer programs.
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Subject Category: 39
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Report Date: May 1997
No. Pages: 284
Funding Organization: WU-953-3600-00
Keywords: Aeroelasticity; Aeroservoelasticity; Computational fluid dynamics; Finite elements; Graphics; Nonlinear aeroelasticity; Structural analysis
Notes: n.a.
- INLET DISTORTION FOR AN F/A-18A AIRCRAFT DURING STEADY AERODYNAMIC CONDITIONS UP TO 60 DEGREE ANGLE OF ATTACK , Technical Memorandum
Authors: Kevin R. Walsh , Andrew J. Yuhas , John G. Williams and William G. Steenken
Report Number: NASA-TM-104329
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This paper
addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from -10 deg to 60 deg angle of attack and from -8 deg to 11 deg angle of sideslip at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to establish a foundation for understanding inlet characteristics at high angle of attack.
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Subject Category: 07
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Report Date: April 1997
No. Pages: 49
Funding Organization: 529 31 04 00 37 00 F-18
Keywords: FA-18A aircraft; High angle of attack; Inlet distortion; Pressure distributions;
Notes: Presented at the High-Angle-of-Attack Technology Conference, NASA Langley Research Center, Hampton, Virginia, Sept. 1719, 1996.
- EVALUATION OF HIGH-SPEED CIVIL TRANSPORT HANDLING QUALITIES CRITERIA WITH SUPERSONIC FLIGHT DATA , Technical Memorandum
Authors: Timothy H. Cox and Dante W. Jackson
Report Number: NASA-TM-4791
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Most flying qualities criteria have been developed from data in the subsonic flight regime. Unique characteristics of supersonic flight raise questions about whether these criteria successfully extend into the supersonic flight regime. Approximately 25 years ago NASA Dryden Flight Research Center addressed this issue with handling qualities evaluations of the XB-70 and YF-12. Good correlations between some of the classical handling qualities parameters, such as the control anticipation parameter as a function of damping, were discovered. More criteria have been developed since these studies. Some of these more recent criteria are being used in designing the High-Speed Civil Transport (HSCT). A second research study recently addressed this issue through flying qualities evaluations of the SR-71 at Mach 3. The research goal was to extend the high-speed flying qualities experience of large airplanes and to evaluate more recent MIL-STD-1797 criteria
against pilot comments and ratings. Emphasis was placed on evaluating the criteria used for designing the HSCT. XB-70 and YF-12 data from the previous research supplemented the SR-71 data. The results indicate that the criteria used in the HSCT design are conservative and should provide good flying qualities for typical high-speed maneuvering. Additional results show correlation between the ratings and comments and criteria for gradual maneuvering with precision control. Correlation is shown between ratings and comments and an extension of the Neal/Smith criterion using normal acceleration instead of pitch rate.
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Subject Category: 08
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Report Date: April 1997
No. Pages: 30
Funding Organization: WU 529-50-24
Keywords: Bandwidth criteria; Flightpath bandwidth; Flying qualities; High-speed civil transport; Low-order equivalent systems; Neal/Smith criteria; SR-71 aircraft; Supersonic flying qualities
- A SUMMARY OF NUMEROUS STRAIN-GAGE LOAD CALIBRATIONS ON AIRCRAFT WINGS AND TAILS IN A TECHNOLOGICAL FORMAT , Technical Memorandum
Authors: Jerald M. Jenkins and V. Michael DeAngelis
Report Number: NASA-TM-4804
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Fifteen aircraft structures that were calibrated for flight loads using strain gages are examined. The primary purpose of this paper is to document important examples of load calibrations on airplanes during the past four decades. The emphasis is placed on studying the physical procedures of calibrating strain-gaged structures and all the supporting analyses and computational techniques that have been used. The results and experiences obtained from actual data from 14 structures (on 13 airplanes and 1 laboratory test structure) are presented. This group of structures includes fins, tails, and wings with a wide variety of aspect ratios. Straight-wing, swept-wing, and delta-wing configurations are studied. Some of the structures have skin-dominant construction; others are spar-dominant. Anisotropic materials, heat shields, corrugated components, nonorthogonal primary structures, and truss-type structures are particular characteristics that are
included.
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Subject Category: 01
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Report Date: July 1997
No. Pages: 146
Funding Organization: WU 505-63-50
Keywords: Anisotropic structure; Flight load; Isotropic structure; Load calibration; Loads measurement; Strain gage
- NASA HYPERSONIC X-PLANE FLIGHT DEVELOPMENT OF TECHNOLOGIES AND CAPABILITIES FOR THE 21ST CENTURY ACCESS TO SPACE , Conference Paper
Authors: John W. Hicks and Gary Trippensee
Report Number: H-2155
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A new family of NASA experimental aircraft (X-planes) is being developed to uniquely, yet synergistically tackle a wide class of technologies to advance low-cost, efficient access to space for a range of payload classes. This family includes two non-air-breathing rocket-powered concepts, the X-33 and the X-34 aircraft, and two air-breathing vehicle concepts, the scramjet-powered Hyper-X and the rocket-based combined-cycle flight vehicle. This report describes the NASA vision for reliable, reusable, fly-to-orbit spacecraft in relation to the current space shuttle capability. These hypersonic X-plane programs, their objectives, and their status are discussed. The respective technology sets and flight program approaches are compared and contrasted. Additionally, the synergy between these programs to advance the entire technology front in a uniform way is discussed. NASA's view of the value of in-flight hypersonic experimentation and technology
development to act as the ultimate crucible for proving and accelerating technology readiness is provided. Finally, an opinion on end technology products and space access capabilities for the 21st century is offered.
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Report Date: April 1997
No. Pages: 10
Funding Organization: n.a.
Keywords: n.a.
Notes: Presented at Proceedings of AGARD Future Aerospace Technology in Service to the Alliance, April 14-18, 1997, Paris, France.
- AN INLET DISTORTION ASSESSMENT DURING AIRCRAFT DEPARTURES AT HIGH ANGLE OF ATTACK FOR AN F/A-18 AIRCRAFT , Technical Memorandum
Authors: William G. Steenken, John G. Williams, Andrew J. Yuhas and Kevin R. Walsh
Report Number: NASA-TM-104328
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The F404-GE-400-powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the quality of inlet airflow during departed flight maneuvers, that is, during flight outside the normal maneuvering envelope where control surfaces have little or no effectiveness. Six nose-left and six nose-right departures were initiated at Mach numbers between 0.3 and 0.4 at an altitude of 35 kft. The entry yaw rates were approximately 40 to 90 deg/sec. Engine surges were encountered during three of the nose-left and one of the nose-right departures. Time-variant inlet-total-pressure distortion levels at the engine face did not significantly exceed those at maximum angle-of-attack and sideslip maneuvers during controlled flight. Surges caused by inlet distortion levels resulted from a combination of high levels of inlet distortion and rapid changes in aircraft position. These rapid changes indicate a combination of engine support and gyroscopic loads being
applied to the engine structure that impact the aerodynamic stability of the compressor through changes in the rotor-to-case clearances. This document presents the slides from an oral presentation.
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Subject Category: 07
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Report Date: March 1997
No. Pages: 32
Funding Organization: WU 505-68-30
Keywords: Aircraft departures; Engine stability limits; F/18-18A airplane; High angle of attack; Inlet distortion
Notes: W.G. Steenken, J.G. Williams, General Electric Aircraft Engines, Cincinnati, OH. A.J. Yuhas, AS&M, Inc., Edwards, CA. K.R. Walsh, NASA Dryden, Edwards, CA. Oral presentation/NASA High-Angle-of-Attack Conf., NASA Langley, Hampton, VA, Sept. 17-19, 1996.
- WAVELET ANALYSES OF F/A-18 AEROELASTIC AND AEROSERVOELASTIC FLIGHT TEST DATA , Technical Memorandum
Authors: Martin J. Brenner
Report Number: NASA-TM-4793
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Time-frequency signal representations combined with subspace identification methods were used to analyze aeroelastic flight data from the F/A-18 Systems Research Aircraft (SRA) and aeroservoelastic data from the F/A-18 High Alpha Research Vehicle (HARV). The F/A-18 SRA data were produced from a wingtip excitation system that generated linear frequency chirps and logarithmic sweeps. HARV data were acquired from digital Schroeder-phased and sinc pulse excitation signals to actuator commands. Nondilated continuous Morlet wavelets implemented as a filter bank were chosen for the time-frequency analysis to eliminate phase distortion as it occurs with sliding window discrete Fourier transform techniques. Wavelet coefficients were filtered to reduce effects of noise and nonlinear distortions identically in all inputs and outputs. Cleaned reconstructed time domain signals were used to compute improved transfer functions. Time and frequency domain subspace
identification methods were applied to enhanced reconstructed time domain data and improved transfer functions, respectively. Time domain subspace performed poorly, even with the enhanced data, compared with frequency domain techniques. A frequency domain subspace method is shown to produce better results with the data processed using the Morlet time-frequency technique.
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Subject Category: 08
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Report Date: April 1997
No. Pages: 31
Funding Organization: 522-32-34-00-RS-00-000
Keywords: Aeroelasticity; Aeroservoelasticity; Aircraft vibration; Modal stability; Structural dynamics; System identification; Wavelet analysis
Notes: Presented at the 38th Structure, Structural Dynamics and Materials Conference Exhibit, AIAA, Kissimmee, Florida, April 7-10, 1997.
- OVERVIEW OF RECENT FLIGHT FLUTTER TESTING RESEARCH AT NASA DRYDEN , Technical Memorandum
Authors: Martin J. Brenner, Richard C. Lind and David F. Voracek
Report Number: NASA-TM-4792
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: In response to the concerns of the aeroelastic community, NASA Dryden Flight Research Center, Edwards, California, is conducting research into improving the flight flutter (including aeroservoelasticity) test process with more accurate and automated techniques for stability boundary prediction. The important elements of this effort so far include the following: 1) excitation mechanisms for enhanced vibration data to reduce uncertainty levels in stability estimates; 2) investigation of a variety of frequency, time, and wavelet analysis techniques for signal processing, stability estimation, and nonlinear identification; and 3) robust flutter boundary prediction to substantially reduce the test matrix for flutter clearance. These are critical research topics addressing the concerns of a recent AGARD Specialists' Meeting on Advanced Aeroservoelastic Testing and Data Analysis. This paper addresses these items using flight test data from the F/A-18
Systems Research Aircraft and the F/A-18 High Alpha Research Vehicle.
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Subject Category: 05
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Report Date: April 1997
No. Pages: 30
Funding Organization: WU 522-32-34
Keywords: Aeroelasticity; Aeroservoelasticity; Flutter; Model validation; Robust stability; Stability prediction; Structural dynamics; System identification; Wavelet analysis
Notes: Presented as AIAA 97-1023 at the 38th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference and Exhibit, Kissimmee, Florida, April 7-10, 1997.
- A MODEL FOR SPACE SHUTTLE ORBITER TIRE SIDE FORCES BASED ON NASA LANDING SYSTEMS RESEARCH AIRCRAFT TEST RESULTS , Technical Memorandum
Authors: John F. Carter, Christopher J. Nagy and Joseph S. Barnicki
Report Number: NASA-TM-4808
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Forces generated by the Space Shuttle orbiter tire under varying vertical load, slip angle, speed, and surface conditions were measured using the Landing System Research Aircraft (LSRA). Resulting data were used to calculate a mathematical model for predicting tire forces in orbiter simulations. Tire side and drag forces experienced by an orbiter tire are cataloged as a function of vertical load and slip angle. The mathematical model is compared to existing tire force models for the Space Shuttle orbiter. This report describes the LSRA and a typical test sequence. Testing methods, data reduction, and error analysis are presented. The LSRA testing was conducted on concrete and lakebed runways at the Edwards Air Force Flight Test Center and on concrete runways at the Kennedy Space Center (KSC). Wet runway tire force tests were performed on test strips made at the KSC using different surfacing techniques. Data were corrected for ply steer forces and
conicity.
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Subject Category: 05
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Report Date: September 1997
No. Pages: 65
Funding Organization: WU 551-15-01
Keywords: Aircraft tire; CV-990; Landing gear; Landing system; Space shuttle orbiter test facility
- FINITE-ELEMENT ANALYSIS OF A MACH-8 FLIGHT TEST ARTICLE USING NONLINEAR CONTACT ELEMENTS , Technical Memorandum
Authors: W. Lance Richards
Report Number: NASA-TM-4796
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A flight test article, called a glove, is required for a Mach-8 boundary-layer experiment to be conducted on a flight mission of the air-launched Pegasus® space booster. The glove is required to provide a smooth, three-dimensional, structurally stable, aerodynamic surface and includes instrumentation to determine when and where boundary-layer transition occurs during the hypersonic flight trajectory. A restraint mechanism has been invented to attach the glove to the wing of the space booster. The restraint mechanism securely attaches the glove to the wing in directions normal to the wing/glove interface surface, but allows the glove to thermally expand and contract to alleviate stresses in directions parallel to the interface surface. A finite-element analysis has been performed using nonlinear contact elements to model the complex behavior of the sliding restraint mechanism. This paper provides an overview of the glove design and presents
details of the analysis that were essential to demonstrate the flight worthiness of the wing-glove test article. Results show that all glove components are well within the allowable stress and deformation requirements to satisfy the objectives of the flight research experiment.
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Subject Category: 39
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Report Date: June 1997
No. Pages: 17
Funding Organization: WU 529-60-24
Keywords: Computational methods; Finite element analysis; Hypersonic flight research; Nonlinear analysis; Structural design
Notes: Presented at the Contact Mechanics 1997 Conference, Madrid, Spain, July 1-3, 1997.
- DEVELOPMENT AND FLIGHT TEST OF AN EMERGENCY FLIGHT CONTROL SYSTEM USING ONLY ENGINE THRUST ON AN MD-11 TRANSPORT AIRPLANE , Technical Paper
Authors: Frank W. Burcham, Jr., John J. Burken, Trindel A. Maine and C. Gordon Fullerton
Report Number: NASA-TP-206217
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An emergency flight control system that uses only engine thrust, called the propulsion-controlled aircraft (PCA) system, was developed and flight tested on an MD-11 airplane. The PCA system is a thrust-only control system, which augments pilot flightpath and track commands with aircraft feedback parameters to control engine thrust. The PCA system was implemented on the MD-11 airplane using only software modifications to existing computers. Results of a 25-hr flight test show that the PCA system can be used to fly to an airport and safely land a transport airplane with an inoperative flight control system. In up-and-away operation, the PCA system served as an acceptable autopilot capable of extended flight over a range of speeds, altitudes, and configurations. PCA approaches, go-arounds, and three landings without the use of any normal flight controls were demonstrated, including ILS-coupled hands-off landings. PCA operation was used to recover from
an upset condition. The PCA system was also tested at altitude with all three hydraulic systems turned off. This paper reviews the principles of throttles-only flight control, a history of accidents or incidents in which some or all flight controls were lost, the MD-11 airplane and its systems, PCA system development, operation, flight testing, and pilot comments.
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Subject Category: 08
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Report Date: November 1997
No. Pages: 91
Funding Organization: WU 522 15 34
Keywords: Emergency Control; MD-11; Propulsion Control; Throttle-only control
- INLET DISTORTION FOR AN F/A-18A AIRCRAFT DURING STEADY AERODYNAMIC CONDITIONS UP TO 60* ANGLE OF ATTACK , Technical Memorandum
Authors: Kevin R. Walsh, Andrew J. Yuhas, John G. Williams, William G. and Steenken
Report Number: NASA-TM-104329
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This paper
addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from -10 degree to 60 degree angle of attack and from -8 degree to 11 degree angle of sideslip at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to establish a foundation for understanding inlet characteristics at high angle of attack.
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Subject Category: 07
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Report Date: April 1997
No. Pages: 50
Funding Organization: 529 31 04 00 37 00 F-18
Keywords: FA-18A aircraft; High angle of attack; Inlet distortion; Pressure distributions;
Notes: Presented at the High-Angle-of-Attack Technology Conference, NASA Langley Research Center, Hampton, VA, Sept. 17-19, 1996.
- A CORRELATION BETWEEN FLIGHT-DETERMINED DERIVATIVES AND WIND-TUNNEL DATA FOR THE X-24B RESEARCH AIRCRAFT , Technical Memorandum
Authors: Alex G. Sim
Report Number: NASA-TM-113084
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Longitudinal and lateral-directional estimates of the aerodynamic derivatives of the X-24B research aircraft were obtained from flight data by using a modified maximum likelihood estimation method. Data were obtained over a Mach number range from 0.35 to 1.72 and over an angle of attack range from 3.5 deg to 15.7 deg. Data are presented for a subsonic and transonic configuration. The flight derivatives were generally consistent and documented the aircraft well. The correlation between the flight data and wind-tunnel predictions is presented and discussed.
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Subject Category: 08
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Report Date: August 1997
No. Pages: 67
Funding Organization: WU 521-71-01
Keywords: Aerodynamic derivatives; Flight-determined X-24B derivatives; Parameter identification; X-24B research aircraft
Notes: Previously published as NASA Service Technical Memorandum SX-3371, March 1976 with the same title.
- FLIGHT-DETERMINED SUBSONIC LONGITUDINAL STABILITY AND CONTROL DERIVATIVES OF THEF-18 HIGH ANGLE OF ATTACK RESEARCH VEHICLE (HARV) WITH THRUST VECTORING , Technical Paper
Authors: Kenneth W. Iliff and Kon-Sheng Charles Wang
Report Number: NASA-TP-206539
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The subsonic longitudinal stability and control derivatives of the F-18 High Angle of Attack Research Vehicle (HARV) are extracted from dynamic flight data using a maximum likelihood parameter identification technique. The technique uses the linearized aircraft equations of motion in their continuous/discrete form and accounts for state and measurement noise as well as thrust-vectoring effects. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics over the aircraft, particularly at high angles of attack. Thrust vectoring was implemented using electrohydraulically-actuated nozzle postexit vanes and a specialized research flight control system. During maneuvers, a control system feature provided independent aerodynamic control surface inputs and independent thrust-vectoring vane inputs, thereby eliminating correlations between the aircraft states and controls. Substantial variations in control excitation and
dynamic response were exhibited for maneuvers conducted at different angles of attack. Opposing vane interactions caused most thrust-vectoring inputs to experience some exhaust plume interference and thus reduced effectiveness. The estimated stability and control derivatives are plotted, and a discussion relates them to predicted values and maneuver quality.
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Subject Category: 08
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Report Date: December 1997
No. Pages: 72
Funding Organization: WU 505 68 50 00 R R 00 000
Keywords: Aerodynamics; Flight-to-wind-tunnel comparisons; F-18 Aircraft; High angle of attack;
- EMERGENCY FLIGHT CONTROL USING ONLY ENGINE THRUST AND LATERAL CENTER-OF-GRAVITY OFFSET: A FIRST LOOK , Technical Memorandum
Authors: Frank W. Burcham, Jr., John Burken, Trindel A. Main and John Bull
Report Number: NASA-TM-4798
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Normally, the damage that results in a total loss of the primary flight controls of a jet transport airplane, including all engines on one side, would be catastrophic. In response, NASA Dryden has conceived an emergency flight control system that uses only the thrust of a wing-mounted engine along with a lateral center-of-gravity (CGY) offset from fuel transfer. Initial analysis and simulation studies indicate that such a system works, and recent high-fidelity simulation tests on the MD-11 and B-747 suggest that the system provides enough control for a survivable landing. This paper discusses principles of flight control using only a wing engine thrust and CGY offset, along with the amount of CGY offset capability of some transport airplanes. The paper also presents simulation results of the throttle-only control capability and closed-loop control of ground track using computer-controlled thrust.
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Subject Category: 08
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Report Date: July 1997
No. Pages: 20
Funding Organization: 522-15-34-00-39-00-IDA
Keywords: B-747; Emergency control; Engine-out control; Lateral center-of-gravity offset; MD-11; Propulsion control
Notes: Published in proceeding for 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, July 6-9, 1997, Seattle, Washington as AIAA-97-3189. Frank W. Burcham, Jr., John Burken, and Trindel A. Main, NASA Dryden Flight Research Center, Edwards, Californi
- DYNAMIC GROUND EFFECT FOR A CRANKED ARROW WING AIRPLANE , Technical Memorandum
Authors: Robert E. Curry
Report Number: NASA-TM-4799
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight-determined ground effect characteristics for an F-16XL airplane are presented and correlated with wind tunnel predictions and similar flight results from other aircraft. Maneuvers were conducted at a variety of flightpath angles. Conventional ground effect flight test methods were used, with the exception that space positioning data were obtained using the differential global positioning system (DGPS). Accuracy of the DGPS was similar to that of optical tracking methods, but it was operationally more attractive. The dynamic flight-determined lift and drag coefficient increments were measurably lower than steady-state wind-tunnel predictions. This relationship is consistent with the results of other aircraft for which similar data are available. Trends in the flight measured lift increments caused by ground effect as a function of flightpath angle were evident but weakly correlated. An engineering model of dynamic ground effect was developed
based on linear aerodynamic theory and super-positioning of flows. This model was applied to the F-16XL data set and to previously published data for an F-15 airplane. In both cases, the model provided an engineering estimate of the ratio between the steady-state and dynamic data sets.
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Subject Category: 02
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Report Date: August 1997
No. Pages: 20
Funding Organization: 529-31-24-00-24-00-RPT
Keywords: Aerodynamic ground effect; Cranked arrow wing; Dynamic ground effect; Flight testing; Supersonic transport
Notes: Presented at AIAA Atmospheric Flight Mechanics Conference, New Orleans, Louisiana, August 11-13, 1997 as AIAA-97-3649.
- SUPERSONIC FLYING QUALITIES EXPERIENCE USING THE SR-71 , Technical Memorandum
Authors: Timothy H. Cox and Dante Jackson
Report Number: NASA-TM-4800
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Approximately 25 years ago NASA Dryden Flight Research Center, Edwards, California, initiated the evaluation of supersonic handling qualities issues using the XB-70 and the YF-12. Comparison of pilot comments and ratings with some of the classical handling qualities criteria for transport aircraft provided information on the usefulness of these criteria and insight into supersonic flying qualities issues. A second research study has recently been completed which again addressed supersonic flying qualities issues through evaluations of the SR-71 in flight at Mach 3. Additional insight into supersonic flying qualities issues was obtained through pilot ratings and comments. These ratings were compared with existing military specifications and proposed criteria for the High Speed Civil Transport. This paper investigates the disparity between pilot comments and the Neal/Smith criteria through a modification of the technique using vertical speed at the
pilot station. The paper specifically addresses the pilot ability to control flightpath and pitch attitude in supersonic flight and pilot displays typical of supersonic maneuvering.
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Subject Category: 08
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Report Date: August 1997
No. Pages: 14
Funding Organization: WU 529-50-24
Keywords: Bandwidth criteria; Flightpath bandwidth; Flying qualities; High Speed Civil
Notes: Presented at the AIAA Atmospheric Flight Mechanics Conference New Orleans, Louisiana, August 11-13, 1997, AIAA 97-3654
- PREDICTED PERFORMANCE OF A THRUST-ENHANCED SR-71 AIRCRAFT WITH AN EXTERNAL PAYLOAD , Technical Memorandum
Authors: Timothy R. Conners
Report Number: NASA-TM-104330
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: NASA Dryden Flight Research Center has completed a preliminary performance analysis of the SR-71 aircraft for use as a launch platform for high-speed research vehicles and for carrying captive experimental packages to high altitude and Mach number conditions. Externally mounted research platforms can significantly increase drag, limiting test time and, in extreme cases, prohibiting penetration through the high-drag, transonic flight regime. To provide supplemental SR-71 acceleration, methods have been developed that could increase the thrust of the J58 turbojet engines. These methods include temperature and speed increases and augmentor nitrous oxide injection. The thrust-enhanced engines would allow the SR-71 aircraft to carry higher drag research platforms than it could without enhancement. This paper presents predicted SR-71 performance with and without enhanced engines. A modified climb-dive technique is shown to reduce fuel consumption when
flying through the transonic flight regime with a large external payload. Estimates are included of the maximum platform drag profiles with which the aircraft could still complete a high-speed research mission. In this case, enhancement was found to increase the SR-71 payload drag capability by 25 percent. The thrust enhancement techniques and performance prediction methodology are described.
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Subject Category: 05
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Report Date: June 1997
No. Pages: 16
Funding Organization: WU 242 33 02 00 23 00 T15
Keywords: Aircraft performance; J58 turbojet engine; Optimal flight profile; SR-71 aircraft; Thrust enhancement; Transonic drag rise
Notes: Presented at the International Gas Turbine and Aeroengine Congress and Exposition, Houston, Texas, June 5-8, 1995.
- A LIMITED IN-FLIGHT EVALUATION OF THE CONSTANT CURRENT LOOP STRAIN MEASUREMENT METHOD , Technical Memorandum
Authors: Candida D. Olney and Joseph V. Collura
Report Number: NASA-TM-104331
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: For many years, the Wheatstone bridge has been used successfully to measure electrical resistance and changes in that resistance. However, the inherent problem of varying lead wire resistance can cause errors when the Wheatstone bridge is used to measure strain in a flight environment. The constant current loop signal-conditioning card was developed to overcome that difficulty. This paper describes a limited evaluation of the constant current loop strain measurement method as used in the F-16XL ship 2 Supersonic Laminar Flow Control flight project. Several identical strain gages were installed in close proximity on a shock fence which was mounted under the left wing of the F-16XL ship 2. Two strain gage bridges were configured using the constant current loop, and two were configured using the Wheatstone bridge circuitry. Flight data comparing the output from the constant current loop configured gages to that of the Wheatstone bridges with respect to
signal output, error, and noise are given. Results indicate that the constant current loop strain measurement method enables an increased output, unaffected by lead wire resistance variations, to be obtained from strain gages.
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Subject Category: 35
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Report Date: August 1997
No. Pages: 32
Funding Organization: WU 529-31-24
Keywords: Circuit theory; Constant current loop; F-16XL aircraft; Strain gage; Structural testings; Wheatstone bridge
Notes: Presented at the Society of Flight Test Engineers Conference, Orlando, Florida, August 18-22, 1997.
- THE USE OF A LIDAR FORWARD-LOOKING TURBULENCE SENSOR FOR MIXED-COMPRESSION INLET UNSTART AVOIDANCE AND GROSS WEIGHT REDUCTION ON A HIGH SPEED CIVILTRANSPORT , Technical Memorandum
Authors: David Soreide , Rodney K. Bogue , Jonathan Seidel and L.J. Ehernberger
Report Number: NASA-TM-104332
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Inlet unstart causes a disturbance akin to severe turbulence for a supersonic commercial airplane. Consequently, the current goal for the frequency of unstarts is a few times per fleet lifetime. For a mixed-compression inlet, there is a tradeoff between propulsion system efficiency and unstart margin. As the unstart margin decreases, propulsion system efficiency increases, but so does the unstart rate. This paper intends to first, quantify that tradeoff for the High Speed Civil Transport (HSCT) and second, to examine the benefits of using a sensor to detect turbulence ahead of the airplane. When the presence of turbulence is known with sufficient lead time to allow the propulsion system to adjust the unstart margin, then inlet unstarts can be minimized while overall efficiency is maximized. The NASA Airborne Coherent Lidar for Advanced In-Flight Measurements program is developing a lidar system to serve as a prototype of the forward-looking sensor.
This paper reports on the progress of this development program and its application to the prevention of inlet unstart in a mixed-compression supersonic inlet. Quantified benefits include significantly reduced takeoff gross weight (TOGW), which could increase payload, reduce direct operating costs, or increase range for the HSCT.
Distribution/Availability: Unclassified - Unlimited
Subject Category: 07
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Report Date: July 1997
No. Pages: 16
Funding Organization: WU 529-50-24-00-RR-00-000
Keywords: SR-71 aircraft; DC-8 aircraft; Inlet aircraft; Alarm systems; Aircraft instrumentation;
Notes: Presented at the 33rd AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, July 6-9, 1997, Seattle, Washington
- STRESS ANALYSIS OF B-52 PYLON HOOKS FOR CARRYING THE X-38 DROP TEST VEHICLE , Technical Memorandum
Authors: William L. Ko
Report Number: TM-97-206218
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The finite-element structural analysis method was used in a two-dimensional stress concentration analysis of a new pylon hook for carrying the X-38 lifting body atmospheric drop test vehicle on the B-52 carrier aircraft. The stress distributions in the hook were obtained, and the critical stress points were identified. The functional relationships between the applied hook load and the induced maximum tangential and shear stresses were established for setting the limit hook load for test flights. By properly representing the X-38 hook with an equivalent curved beam, the conventional curved beam theory predicted the values of the maximum tangential and shear stresses quite close to those calculated from the finite-element analysis. The equivalent curved beam method may be of practical value during the initial stage of the hook design in estimating the critical stresses and failure loads with reasonable accuracy.
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Subject Category: 39
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Report Date: October 1997
No. Pages: 33
Funding Organization: WU 250 88 00
Keywords: Equivalent curved beam theory; Finite element stress analysis ; Stress concentrations; X-38 hooks
- A TECHNIQUE FOR TRANSIENT THERMAL TESTING OF THICK STRUCTURES , Technical Memorandum
Authors: Thomas J. Horn, W. Lance Richards and Leslie Gong
Report Number: NASA-TM-4803
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A new open-loop heat flux control technique has been developed to conduct transient thermal testing of thick, thermally-conductive aerospace structures. This technique uses calibration of the radiant heater system power level as a function of heat flux, predicted aerodynamic heat flux, and the properties of an instrumented test article. An iterative process was used to generate open-loop heater power profiles prior to each transient thermal test. Differences between the measured and predicted surface temperatures were used to refine the heater power level command profiles through the iteration process. This iteration process has reduced the effects of environmental and test system design factors, which are normally compensated for by closed-loop temperature control, to acceptable levels. The final revised heater power profiles resulted in measured temperature time histories which deviated less than 25 deg F from the predicted surface temperatures.
Distribution/Availability: Unclassified - Unlimited
Subject Category: 09
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Report Date: July 1997
No. Pages: 22
Funding Organization: WU 529-60-24-00-17-00-DOC
Keywords: Heat flux; Heat transfer; Radiant heating; Structural tests; Thermal testing; Transient thermal simulation
Notes: Presented at the International Society for Optical Engineers Symposium, San Diego, California, July 27-August 1, 1997.
- COMPARATIVE OPTICAL MEASUREMENTS OF AIRSPEED AND AEROSOLS ON A DC-8 AIRCRAFT , Technical Memorandum
Authors: Rodney Bogue, Rick McGann, Thomas Wagener, John Abbiss and Anthony Smart
Report Number: NASA-TM-113083
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: NASA Dryden supported a cooperative flight test program on the NASA DC-8 aircraft in November 1993. This program evaluated optical airspeed and aerosol measurement techniques. Three brassboard optical systems were tested. Two were laser Doppler systems designed to measure free-stream-referenced airspeed. The third system was designed to characterize the natural aerosol statistics and airspeed. These systems relied on optical backscatter from natural aerosols for operation. The DC-8 aircraft carried instrumentation that provided real-time flight situation information and reference data on the aerosol environment. This test is believed to be the first to include multiple optical airspeed systems on the same carrier aircraft, so performance could be directly compared. During 23 hr of flight, a broad range of atmospheric conditions was encountered, including aerosol-rich layers, visible clouds, and unusually clean (aerosol-poor) regions. Substantial
amounts of data were obtained. Important insights regarding the use of laser-based systems of this type in an aircraft environment were gained. This paper describes the sensors used and flight operations conducted to support the experiments. The paper also briefly describes the general results of the experiments.
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Subject Category: 06
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Report Date: July 1997
No. Pages: 28
Funding Organization: WU 529 50 24 00 RR 00 000
Keywords: Aerosols; Aircraft instrumentation; Airspeed indicators; Laser optical radar
Notes: Presented at the 16th International Congress on Instrumentation in Aerospace Simulation Facilities, Wright-Patterson AFB, Dayton, Ohio, July 17-21, 1995.
- FLIGHT DEMONSTRATION OF A SHOCK LOCATION SENSOR USING CONSTANT VOLTAGE HOT-FILM ANEMOMETRY , Technical Memorandum
Authors: Timothy R. Moes, Garimella R. Sarma and Siva M. Mangalam
Report Number: NASA-TM-4806
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight tests have demonstrated the effectiveness of an array of hot-film sensors using constant voltage anemometry to determine shock position on a wing or aircraft surface at transonic speeds. Flights were conducted at the NASA Dryden Flight Research Center using the F-15B aircraft and Flight Test Fixture (FTF). A modified NACA 0021 airfoil was attached to the side of the FTF, and its upper surface was instrumented to correlate shock position with pressure and hot-film sensors. In the vicinity of the shock-induced pressure rise, test results consistently showed the presence of a minimum voltage in the hot-film anemometer outputs. Comparing these results with previous investigations indicate that hot-film anemometry can identify the location of the shock-induced boundary layer separation. The flow separation occurred slightly forward of the shock-induced pressure rise for a laminar boundary layer and slightly aft of the start of the pressure rise
when the boundary layer was tripped near the airfoil leading edge. Both minimum mean output and phase reversal analyses were used to identify the shock location.
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Subject Category: 06
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Report Date: August 1997
No. Pages: 27
Funding Organization: WU 529 21 34 00 38 00 F15
Keywords: Transonic Shocks; Shock Impingement; Boundary Layers; Hot-Films; Anemometry; Constant Voltage Anemometry
Notes: Presented at the 1997 Society of Flight Test Engineers Symposium, Orlando, Florida, August 18-22, 1997. Timothy R. Moes, Dryden Flight Research Center, Edwards, California; Garimella R. Sarma, and Siva M. Mangalam, Tao of Systems Integration, Inc., Hampto
- IN-FLIGHT TRANSPORT PERFORMANCE OPTIMIZATION: AN EXPERIMENTAL FLIGHT RESEARCH PROGRAM AND AN OPERATIONAL SCENARIO , Technical Memorandum
Authors: Glenn Gilyard
Report Number: NASA-TM-206229
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A flight research program exploring the practical application of real-time performance optimization based on aircraft measurements and calculation of incremental drag from forced-response maneuvers is presented. The outboard ailerons of the L-1011 test bed aircraft were modified to provide for symmetric deflections to permit a recambering of the wing in that localized area, which in turn modifies the entire wing load distribution. The National Aeronautics and Space Administration developed an onboard research engineering test station from which the flight experiments are conducted and all analyses, both qualitative and quantitative, are performed in a real-time or near real-time manner. Initial flight test results are presented that indicate real-time drag minimization is attainable. An approach to an operational implementation of adaptive performance optimization on current and future commercial and military transports is discussed with the goal of
keeping the required modifications simple and the pilot interface minimal and user friendly.
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Subject Category: 02-01, 02-03, and 03-01
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Report Date: October 1997
No. Pages: 22
Funding Organization: 522-15-34-00-39-00-210
Keywords: Aircraft Performance; Drag Reduction; Fuel Consumption; Wing Camber; Flight Optimization; and Optimization
Notes: Presented at the 16th Digital Avionics Systems Conference, Irvine, California, October 26-30, 1997
- PRODUCTION SUPPORT FLIGHT CONTROL COMPUTERS: RESEARCH CAPABILITY FOR F/A-18 AIRCRAFT AT DRYDEN FLIGHT RESEARCH CENTER , Technical Memorandum
Authors: John Carter
Report Number: NASA-TM-206233
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: NASA Dryden Flight Research Center (DFRC) is working with the United States Navy to complete ground testing and initiate flight testing of a modified set of F/A-18 flight control computers. The Production Support Flight Control Computers (PSFCC) can give any fleet F/A-18 airplane an in-flight, pilot-selectable research control law capability. NASA DFRC can efficiently flight test the PSFCC for the following four reasons: (1) Six F/A-18 chase aircraft are available which could be used with the PSFCC. (2) An F/A-18 processor-in-the-loop simulation exists for validation testing. (3) The expertise has been developed in programming the research processor in the PSFCC. (4) A well-defined process has been established for clearing flight control research projects for flight. This report presents a functional description of the PSFCC. Descriptions of the NASA DFRC facilities, PSFCC verification and validation process, and planned PSFCC projects are also
provided.
Distribution/Availability: Unclassified - Unlimited
Subject Category: 09
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Report Date: October 1997
No. Pages: 20
Funding Organization: 242-33-02-00-25-00-000
Keywords: Flight control computers; Control laws; F/A-18 Aircraft; Production support flight
Notes: Presented at the 16th Digital Avionics Systems Conference, Irvine, California, October 26-30, 1997
- LINEARIZED POSTSTALL AERODYNAMIC AND CONTROL LAW MODELS OF THE X-31A AIRCRAFT AND COMPARISON WITH FLIGHT DATA , Technical Memorandum
Authors: Patrick C. Stoliker, John T. Bosworth and Jennifer Georgie
Report Number: NASA-TM-97-206318
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The X-31A aircraft has a unique configuration that uses thrust-vector vanes and aerodynamic control effectors to provide an operating envelope to a maximum 70 deg angle of attack, an inherently nonlinear portion of the flight envelope. This report presents linearized versions of the X-31A longitudinal and lateral-directional control systems, with aerodynamic models sufficient to evaluate characteristics in the poststall envelope at 30 deg, 45 deg, and 60 deg angle of attack. The models are presented with detail sufficient to allow the reader to reproduce the linear results or perform independent control studies. Comparisons between the responses of the linear models and flight data are presented in the time and frequency domains to demonstrate the strengths and weaknesses of the ability to predict high-angle-of-attack flight dynamics using linear models. The X-31A six-degree-of-freedom simulation contains a program that calculates linear
perturbation models throughout the X-31A flight envelope. The models include aerodynamics and flight control system dynamics that are used for stability, controllability, and handling qualities analysis. The models presented in this report demonstrate the ability to provide reasonable linear representations in the poststall flight regime.
Distribution/Availability: Unclassified - Unlimited
Subject Category: 08
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Report Date: December 1997
No. Pages: 66
Funding Organization: WU 529-30-04
Keywords: Aircraft flight control system; Linear aerodynamics model; Rigid-body dynamic model; State space; X-31A airplane
Notes: Errata added January 22, 1998, an error in equation (3) and equation (5) on page 9 was corrected.
- TAILORED EXCITATION FOR MULTIVARIABLE STABILITY-MARGIN MEASUREMENT APPLIED TO THE X-31A NONLINEAR SIMULATION , Technical Memorandum
Authors: John T. Bosworth and John J. Burken
Report Number: NASA-TM-113085
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Safety and productivity of the initial flight test phase of a new vehicle have been enhanced by developing the ability to measure the stability margins of the combined control system and vehicle in flight. One shortcoming of performing this analysis is the long duration of the excitation signal required to provide results over a wide frequency range. For flight regimes such as high angle of attack or hypersonic flight, the ability to maintain flight condition for this time duration is difficult. Significantly reducing the required duration of the excitation input is possible by tailoring the input to excite only the frequency range where the lowest stability margin is expected. For a multiple-input/multiple-output system, the inputs can be simultaneously applied to the control effectors by creating each excitation input with a unique set of frequency components. Chirp-Z transformation algorithms can be used to match the analysis of the results to
the specific frequencies used in the excitation input. This report discusses the application of a tailored excitation input to a high-fidelity X-31A linear model and nonlinear simulation. Depending on the frequency range, the results indicate the potential to significantly reduce the time required for stability measurement.
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Report Date: August 1997
No. Pages: 20
Funding Organization: WU-529-30-04
Keywords: Control; Excitation; Fourier; Margin; Stability
Notes: Presented at the Society of Flight Test Engineers Conference, Orlando, Florida, August 18-22, 1997.
- THE WORLD WIDE WEB AS A MEDIUM OF INSTRUCTION: WHAT WORKS AND WHAT DOESN'T , Conference Publication
Authors: Marianne McCarthy, Barbara Grabowski, Angel Hernandez, Tiffany Koszalka and Lee Duke
Report Number: NASA-CP-3358
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A conference was held on March 18-20, 1997 to investigate the lessons learned by the Aeronautics Cooperative Agreement Projects with regard to the most effective strategies for developing instruction for the World Wide Web. The conference was a collaboration among the NASA Aeronautics and Space Transportation Technology Centers (Ames, Dryden, Langley, and Lewis), NASA Headquarters, the University of Idaho and The Pennsylvania State University. The conference consisted of presentations by the Aeronautics Cooperative Agreement Teams, the University of Idaho, and working sessions in which the participants addressed teacher training and support, technology, evaluation and pedagogy. The conference was also undertaken as part of the Dryden Learning Technologies Project which is a collaboration between the Dryden Education Office and The Pennsylvania State University. The DFRC Learning Technology Project goals relevant to the conference are; conducting an
analysis of current teacher needs, classroom infrastructure and exemplary instructional World Wide Web sites, and developing models for Web-enhanced learning environments that optimize teaching practices and student learning.
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Report Date: September 1997
No. Pages: 134
Funding Organization: WU 332-41-00
Keywords: Education; Learning technology; Pedagogy; Teacher training; Technology; World Wide Web
Notes: Conference Proceedings of the March Learning Technologies Conference, NASA Dryden Flight Research Center, Edwards, California. Also Presented at Ames Research Center, Learning Technology Conference, May 1997.
- AN IMPACT-LOCATION ESTIMATION ALGORITHM FOR SUBSONIC UNINHABITED AIRCRAFT , Technical Memorandum
Authors: Jeffrey E. Bauer and Edward Teets
Report Number: NASA-TM-97-206299
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An impact-location estimation algorithm is being used at the NASA Dryden Flight Research Center to support range safety for uninhabited aerial vehicle flight tests. The algorithm computes an impact location based on the descent rate, mass, and altitude of the vehicle and current wind information. The predicted impact location is continuously displayed on the range safety officer's moving map display so that the flightpath of the vehicle can be routed to avoid ground assets if the flight must be terminated. The algorithm easily adapts to different vehicle termination techniques and has been shown to be accurate to the extent required to support range safety for subsonic uninhabited aerial vehicles. This paper describes how the algorithm functions, how the algorithm is used at NASA Dryden, and how various termination techniques are handled by the algorithm. Other approaches to predicting the impact location and the reasons why they were not selected
for real-time implementation are also discussed.
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Report Date: November 1997
No. Pages: 28
Funding Organization: WU 529 10 04 00 29 00
Keywords: Flight termination; Flight testing; Impact prediction; Range safety; Uninhabited aircraft
- FAILURE STUDY OF COMPOSITE MATERIALS BY THE YEH-STRATTON CRITERION , Technical Memorandum
Authors: Hsien-Yang Yeh and W. Lance Richards
Report Number: NASA-TM-113087
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The newly developed Yeh-Stratton (Y-S) Strength Criterion was used to study the failure of composite materials with central holes and normal cracks. To evaluate the interaction parameters for theY-S failure theory, it is necessary to perform several biaxial loading tests. However, it is indisputable that the inhomogeneous and anisotropic nature of composite materials have made their own contribution to the complication of the biaxial testing problem. To avoid the difficulties of performing many biaxial tests and still consider the effects of the interaction term in theY-S Criterion, a simple modification of theY-S Criterion was developed. The preliminary predictions by the modified Y-S Criterion were relatively conservative compared to the testing data. Thus, the modified Y-S Criterion could be used as a design tool. To further understand the composite failure problem, an investigation of the damage zone in front of the crack tip coupled with the
Y-S Criterion is imperative.
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Report Date: October 1997
No. Pages: 9
Funding Organization: WU 529-50-14
Keywords: Composite materials; Composite research; Failure theory; Macromechanics; Stress concentrations
Notes: Dr. Hsien-Yang Yeh's work was conducted in support of NASA Contract NCC 4-107.
- ON-LINE MODAL STATE MONITORING OF SLOWLY TIME-VARYING STRUCTURES , Contractor Report
Authors: Erik A. Johnson, Lawrence A. Bergman and Petros G. Voulgaris
Report Number: NASA-CR-198057
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Monitoring the dynamic response of structures is often performed for a variety of reasons. These reasons include condition-based maintenance, health monitoring, performance improvements, and control. In many cases the data analysis that is performed is part of a repetitive decision-making process, and in these cases the development of effective on-line monitoring schemes help to speed the decision-making process and reduce the risk of erroneous decisions. This report investigates the use of spatial modal filters for tracking the dynamics of slowly time-varying linear structures. The report includes an overview of modal filter theory followed by an overview of several structural system identification methods. Included in this discussion and comparison are Hinfinity, eigensystem realization, and several time-domain least squares approaches. Finally, a two-stage adaptive on-line monitoring scheme is developed and evaluated.
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Report Date: October 1997
No. Pages: 188
Funding Organization: 529-50-04-00-RR-00-000
Keywords: Modal filtering; Reciprocal modal vector; Parameter estimation; Structural dynamics; On-line health monitoring
Notes: Technical Monitor: Lawrence C. Freudinger, NASA Dryden. NASA Contract NAG 2-4001.
- FLIGHT TESTING THE X-36—THE TEST PILOT'S PERSPECTIVE , Contractor Report
Authors: Laurence A. Walker
Report Number: NASA-CR-198058
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
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Report Date: October 1997
No. Pages: 15
Funding Organization: WU 529 30 04
Keywords: Agility; High angle of attack; RPV; Tailless; Thrust vectoring; X-36 remotely piloted vehicle
Notes: Performing Organzation: Boeing AS&T Phantom Works, St. Louis Mo. Tech. Monitors: M. Sumich (Ames Research Center) and D. Bessette (NASA Dryden). Presented at Society of Experimental Test Pilots Symposium, Beverly Hills, CA, Sept. 25-27, 1997.
- FIBER OPTIC EXPERIENCE WITH THE SMART ACTUATION SYSTEM ON THE F-18 SYSTEMS RESEARCH AIRCRAFT , Technical Memorandum
Authors: Eddie Zavala
Report Number: NASA-TM-206223
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: High bandwidth, immunity to electromagnetic interference, and potential weight savings have led to the development of fiber optic technology for future aerospace vehicle systems. This technology has been incorporated in a new smart actuator as the primary communication interface. The use of fiber optics simplified system integration and significantly reduced wire count. Flight test results showed that fiber optics could be used in aircraft systems and identified critical areas of development of fly-by-light technology. This paper documents the fiber optic experience gained as a result of this program, and identifies general design considerations that could be used in a variety of specific applications of fiber optic technology. Environmental sensitivities of fiber optic system components that significantly contribute to optical power variation are discussed. Although a calibration procedure successfully minimized the effect of fiber optic
sensitivities, more standardized calibration methods are needed to ensure system operation and reliability in future aerospace vehicle systems.
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Report Date: October 1997
No. Pages: 22
Funding Organization: 529-31-14-00-36-00-SRA
Keywords: Fiber optics; smart actuator; F-18 Systems Research Aircraft; electrical actuation
Notes: Presented at the 16th Digital Avionics Systems Conference, Irvine, California, October 26-30, 1997
- A PRESENTATION ON ROBUST FLUTTER MARGIN ANALYSIS AND A FLUTTEROMETER , Technical Memorandum
Authors: Rick C. Lind
Report Number: NASA-TM-97-206220
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This paper documents an invited presentation given to The Boeing Company, Seattle, Washington, on September 9, 1997. The audience consisted of structural dynamic and flight test engineers from the Boeing Commercial Airplane Group who were interested in discussing research which may be applied to future flight flutter test programs. A method to compute robust flutter margins is described which is a significant departure from traditional methods. This method uses the structured singular value, mu, to compute a flutter margin which directly accounts for modeling errors such that a worst-case flutter margin is computed with respect to those errors. This method may be applied in several ways. A post-flight application uses data sets from multiple test points to compute worst-case flutter margins and a worst-case flight envelope. An on-line implementation computes flutter margins at each test point to track the flutter margins during a flight test. This
on-line implementation is the basis for a flutterometer flight test tool that displays the distance to flutter at a given test point. Such a tool was not previously possible using traditional flutter flight test analysis methods. The F/A-18 System Research Aircraft was used to demonstrate these applications using flight data recorded from test points throughout the flight envelope.
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Report Date: October 1997
No. Pages: 65
Funding Organization: WU 529-31-14
Keywords: Aeroelasticity; Flight test; Flutter; Robust stability; Structured singular value
Notes: Oral presentation given at The Boeing Company, Seattle, Washington, on September 9, 1997. Rick C. Lind is a National Research Council research associate.
- PERFORMANCE OF AN ELECTRO-HYDROSTATIC ACTUATOR ON THE F-18 SYSTEMS RESEARCH AIRCRAFT , Technical Memorandum
Authors: Robert Navarro
Report Number: NASA-TM-206224
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An electro-hydrostatic actuator was evaluated at NASA Dryden Flight Research Center, Edwards, California. The primary goal of testing this actuator system was the flight demonstration of power-by-wire technology on a primary flight control surface. The electro-hydrostatic actuator uses an electric motor to drive a hydraulic pump and relies on local hydraulics for force transmission. This actuator replaced the F-18 standard left aileron actuator on the F-18 Systems Research Aircraft and was evaluated throughout the Systems Research Aircraft flight envelope. As of July 24, 1997 the electro-hydrostatic actuator had accumulated 23.5 hours of flight time. This paper presents the electro-hydrostatic actuator system configuration and component description, ground and flight test plans, ground and flight test results, and lessons learned. This actuator performs as well as the standard actuator and has more load capability than required by aileron actuator
specifications of McDonnell-Douglas Aircraft, St. Louis, Missouri. The electro-hydrostatic actuator system passed all of its ground tests with the exception of one power-off test during unloaded dynamic cycling.
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Report Date: October 1997
No. Pages: 37
Funding Organization: 529-31-14-00-36-00-SRA
Keywords: Electro-hydrostatic actuator; F-18 Systems Research Aircraft; Performance; Ground test; Flight test
Notes: Presented at the 16th Digital Avionics Systems Conference, Irvine, California, October 26-30, 1997
- UTILIZING FLIGHT DATA TO UPDATE AEROELASTIC STABILITY ESTIMATES , AIAA Conference Paper
Authors: Rick Lind and Marty Brenner
Report Number: H-2213
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Stability analysis of high performance aircraft must account for errors in the system model. A method for computing flutter margins that incorporates flight data has been developed using robust stability theory. This paper considers applying this method to update flutter margins during a post-flight or on-line analysis. Areas of modeling uncertainty that arise when using flight data with this method are investigated. The amount of conservatism in the resulting flutter margins depends on the flight data sets used to update the model. Post-flight updates of flutter margins for an F/A-18 are presented along with a simulation of on-line updates during a flight test.
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Report Date: August 1997
No. Pages: n.a.
Funding Organization: 505 63 50
Keywords: Aeroelasticity; Flight data; Robust stability; Structured singular value
Notes: Presented at AIAA Atmospheric Flight Mechanics Conference, New Orleans, Louisiana, August 1997.
- ESTIMATION OF MODAL PARAMETERS USING A WAVELET-BASED APPROACH , Technical Memorandum
Authors: Rick Lind, Marty Brenner and Sydney M. Haley
Report Number: NASA-TM-97-206300
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Modal stability parameters are extracted directly from aeroservoelastic flight test data by decomposition of accelerometer response signals into time-frequency atoms. Logarithmic sweeps and sinusoidal pulses are used to generate DAST closed loop excitation data. Novel wavelets constructed to extract modal damping and frequency explicitly from the data are introduced. The so-called Haley and Laplace wavelets are used to track time-varying modal damping and frequency in a matching pursuit algorithm. Estimation of the trend to aeroservoelastic instability is demonstrated successfully from analysis of the DAST data.
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Report Date: November 1997
No. Pages: 14
Funding Organization: 505 63 50
Keywords: Aeroelasticity; Wavelet; Modal; Parameter estimation; System identification
Notes: Presented at AIAA Atmospheric Flight Mechanics Conference, New Orleans, Louisiana, August 8-11, 1997
- CHOOSING SENSOR CONFIGURATION FOR A FLEXIBLE STRUCTURE USING FULL CONTROL , AIAA Conference Paper
Authors: Rick Lind, Volkan Nalbantogul and Gary Balas
Report Number: AIAA-97-3745
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Optimal locations and types for feedback sensors which meet design constraints and control requirements are difficult to determine. This paper introduces an approach to choosing a sensor configuration based on Full Control synthesis. A globally optimal Full Control compensator is computed for each member of a set of sensor configurations which are feasible for the plant. The sensor configuration associated with the Full Control system achieving the best closed-loop performance is chosen for feedback measurements to an output feedback controller. A flexible structure is used as an example to demonstrate this procedure. Experimental results show sensor configurations chosen to optimize the Full Control performance are effective for output feedback controllers.
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Report Date: August 1997
No. Pages: 11
Funding Organization: 505-63-50
Keywords: Flexible structure; Full control; Linear matrix inequality; Sensor placement; Structured singular value
Notes: Presented at the AIAA Guidance and Control Conference, New Orleans, Louisiana, August 1997
- IMPROVED FLIGHT TEST PROCEDURES FOR FLUTTER CLEARANCE , Conference Paper
Authors: Rick Lind, Martin J. Brenner and Lawrence C. Freudinger
Report Number: H-2225
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Flight flutter testing is an integral part of flight envelope clearance. This paper discusses advancements in several areas that are being investigated to improve efficiency and safety of flight test programs. Results are presented from recent flight testing of the F/A-18 Systems Research Aircraft. A wingtip excitation system was used to generate aeroelastic response data. This system worked well for many flight conditions but still displayed some anomalies. Wavelet processing is used to analyze the flight data. Filtered transfer functions are generated that greatly improve system identification. A flutter margin is formulated that accounts for errors between a model and flight data. Worst-case flutter margins are computed to demonstrate the flutter boundary may lie closer to the flight envelope than previously estimated. This paper concludes with developments for a distributed flight analysis environment and on-line health monitoring.
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Report Date: June 1997
No. Pages: 8
Funding Organization: 505-63-50
Keywords: Aeroelasticity; Flutter; Health monitoring; Robust stability; Structured singular value; Wavelet
Notes: CEAS International Forum on Aerostability and Structural Dynamics, Rome, Italy, June 1997
- A WORST-CASE APPROACH FOR ON-LINE FLUTTER PREDICTION , Conference Paper
Authors: Rick Lind and Marty Brenner
Report Number: H-2226
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Worst-case flutter margins may be computed for a linear model with respect to a set of uncertainty operators using the structured singular value. This paper considers an on-line implementation to compute these robust margins in a flight test program. Uncertainty descriptions are updated at test points to account for unmodeled time-varying dynamics of the airplane by ensuring the robust model is not invalidated by measured flight data. Robust margins computed with respect to this uncertainty remain conservative to the changing dynamics throughout the flight. A simulation clearly demonstrates this method can improve the efficiency of flight testing by accurately predicting the flutter margin to improve safety while reducing the necessary flight time.
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Report Date: June 1997
No. Pages: 8
Funding Organization: 505-63-50
Keywords: Aeroelasticity; Flight test; Flutter; Robust stability; Structural dynamics; Structural singular value
Notes: CEAS International Forum on Aerostability and Structural Dynamics, Rome, Italy, June 1997
- WORST-CASE FLUTTER MARGINS FROM F/A-18 AIRCRAFT AEROELASTIC DATA , AIAA Conference
Authors: Rick Lind and Marty Brenner
Report Number: AIAA-97-1266
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An approach for computing worst-case flutter margins has been formulated in a robust stability framework. Uncertainty operators are included with a linear model to describe modeling errors and flight variations. The structured singular value, mu, computes a stability margin which directly accounts for these uncertainties. This approach introduces a new method of computing flutter margins and an associated new parameter for describing these margins. The mu margins are robust margins which indicate worst-case stability estimates with respect to the defined uncertainty. Worst-case flutter margins are computed for the F/A-18 SRA using uncertainty sets generated by flight data analysis. The robust margins demonstrate flight conditions for flutter may lie closer to the flight envelope than previously estimated by p-k analysis.
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Report Date: April 1997
No. Pages: n.a.
Funding Organization: 529-50-04
Keywords: Flight test; Flutter; Robust stability; Stability margins; Structural dynamics
Notes: Presented at AIAA Structures, Structural Dynamics, and Materials Conference, Orlando, Florida, April 1997.
- SPECTRAL DENSITY OF LASER BEAM SCINTILLATION IN WIND TURBULENCE: I. THEORY , Journal Article
Authors: A. V. Balakrishnan
Report Number: H-2321
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The temporal spectral density of the log-amplitude scintillation of a laser beam wave due to a spatially dependent vector-valued crosswind (deterministic as well as random) is evaluated. The path weighting functions for normalized spectral moments are derived, and offer a potential new technique for estimating the wind velocity profile. The Tatarskii-Klyatskin stochastic propagation equation for the Markov turbulence model is used with the solution approximated by the Rytov method. The Taylor 'frozen-in' hypothesis is assumed for the dependence of the refractive index on the wind velocity, and the Kolmogorov spectral density is used for the refractive index field.
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Report Date: April 1997
No. Pages: 34
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: Computational and Applied Mathematics. To appear 1997. 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.
- THEORETICAL LIMITS OF DAMPING ATTAINABLE BY SMART BEAMS WITH RATE FEEDBACK , Conference Paper
Authors: A. V. Balakrishnan
Report Number: H-2322
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Using a generally accepted model we present a comprehensive analysis (within the page limitation) of an Euler-Bernoulli beam with PZT sensor-actuator and pure rate feedback. The emphasis is on the root locus—the dependence of the attainable damping on the feedback gain. There is a critical value of the gain beyond which the damping decreases to zero. We construct the time-domain response using semigroup theory, and show that the eigenfunctions form a Riesz basis, leading to a 'modal' expansion.
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Report Date: April 1997
No. Pages: 13
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: Proc. of SPIE's 4th Annual Symp. on Smart Structures and Materials, March 2-6, 1997, San Diego, Calfornia. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.
- EXPERIMENTAL INVESTIGATIONS OF JET IMPINGEMENT HEAT TRANSFER USING THERMOCHROMIC LIQUID CRYSTALSBRIAN PAUL DEMPSEY , UCLA Ph.D. dissertation, 1997
Authors: Brian Paul Dempsey
Report Number: H-2324
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Jet impingement cooling of a hypersonic airfoil leading edge is experimentally investigated using thermochromic liquid crystals (TLCs) to measure surface temperature. The experiment uses computer data acquisition with digital imaging of the TLCs to determine heat transfer coefficients during a transient experiment. The data reduction relies on analysis of a coupled transient conduction - convection heat transfer problem that characterizes the experiment. The recovery temperature of the jet is accounted for by running two experiments with different heating rates, thereby generating a second equation that is used to solve for the recovery temperature. The resulting solution requires a complicated numerical iteration that is handled by a computer. Because the computational data reduction method is complex, special attention is paid to error assessment. The error analysis considers random and systematic errors generated by the instrumentation along with
error generated by the approximate nature of the numerical methods. Results of the error analysis show that the experimentally determined heat transfer coefficients are accurate to within 15 percent. The error analysis also shows that the recovery temperature data may be in error by more than 50 percent. The results show that the recovery temperature data is only reliable when the recovery temperature of the jet is greater than 5 degrees Celsius, i.e., the jet velocity is in excess of 100 m/s. Parameters that were investigated include nozzle width, distance from the nozzle exit to the airfoil surface, and jet velocity. Heat transfer data is presented in graphical and tabular forms. An engineering analysis of hypersonic airfoil leading edge cooling is performed using the results from these experiments. Several suggestions for the improvement of the experimental technique are discussed.
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Report Date: June 1997
No. Pages: 163
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: This document is the result of work sponsored by NASA Dryden Flight Research Center and performed through the UCLA, Flight Systems Research Center. The contract monitor was Robert Quinn, and funding was facilitated by Dr. Kenneth Iliff.
- MEMS FOR AERODYNAMIC CONTROL , Conference Paper
Authors: Motoaki Kimura, Steve Tung, Chih-Ming Ho, Fukang Jiang and Yu-Chong Tai
Report Number: H-2325
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Multiple arrays of micromachined micro shear-stress sensors have been designed and fabricated to spatially and temporally resolve the small streamwise streaks in the near-wall region of a turbulent boundary layer. By using these sensors, the instantaneous surface shear-stress distribution of a turbulent boundary layer at high Reynolds numbers has been measured. Based on the measurement, the physical quantities associated with the high shear-stress streaks, such as their length, width and peak shear-stress level, are obtained. We found that a high correlation exists between the peak shear stress level and the front-end shear-stress gradient of a high shear-stress streak. This important property is currently being applied to the design of a real-time flow control logic. The control logic will be used to initiate a micro actuator, which interacts with the detected streak for the purpose of reducing local shear stress. The interaction between an
actuator with near wall vortices was first investigated in a laminar boundary layer. The results show that the level of reduction is related to the product of the actuator's peak amplitude and oscillating frequency. This result suggests the possibility of achieving shear-stress reduction by using high-frequency and low-amplitude actuators.
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Report Date: June 1997
No. Pages: 9
Funding Organization: NASA Grant number NCC-2-374
Keywords: n.a.
Notes: Presented at 28th AIAA Fluid Dynamics Conf., 4th AIAA Shear Flow Control Conf., June-July 1997, Colorado. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.
- HEART FIBRILLATION AND PARALLEL SUPERCOMPUTERS , Conference Paper
Authors: B. Y. Kogan, W. J. Karplus and E. E. Chudin
Report Number: H-2326
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The Luo and Rudy III cardiac cell mathematical model is implemented on the parallel supercomputer CRAY-T3D. The splitting algorithm combined with variable time step and an explicit method of integration provide reasonable solution times and almost perfect scaling for rectilinear wave propagation. The computer simulation makes it possible to observe new phenomena: the break-up of spiral waves caused by intracellular calcium dynamics and the non-uniformity of the calcium distribution in space during the onset of the spiral wave.
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Report Date: June 1997
No. Pages: 10
Funding Organization: NASA Grant number NCC-2-374
Keywords: Heart fibrillation; Intracellular calcium effect; Simulation on parallel computers
Notes: Presented at Int. Conf. on Informatics and Control, Vol. 3, St. Petersburg, Russia, 1997, pp. 857-865. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.
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