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  1. SR-71 EXPERIMENT ON PROPAGATION OF SONIC BOOMS
    Authors: Edward A. Haering, Jr.
    Report Number: DRC-95-32
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight experiment was recently conducted at the NASA's Dryden Flight Research Center in Edwards, California, using an SR-71, an F-16XL, and a YO-3A airplane to study the propagation of sonic booms. This work is geared toward developing a high-speed civil transport (HSCT) aircraft for operational use early in the next century. At supersonic speeds, an aircraft generates numerous shock waves that emanate from such major components as the nose, canopy, inlets, wings, and vertical tails. These multiple shock waves tend to merge into a strong bow shock and a strong tail shock as they propagate through the atmosphere. At present, supercomputers and wind tunnels are used to predict the structures of these shock waves, but only within a few body lengths of the aircraft. Other computational techniques are then used to determine the propagation and merging of these shock waves down to ground level. To verify and enhance the quality of these computational propagation techniques, a database of sonic-boom measurements at various altitudes was gathered for use by the aerospace industry, universities, and NASA research centers. These organizations will use the enhanced computational techniques in the design of the HSCT. Varying the design of a HSCT could help soften the intensity of sonic booms at ground level.
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    HTMLReport Date: January 1996
    No. Pages: 2
    Keywords:      F-16XL; Propagation; Sonic boom; SR-71; YO-3A
    Notes: Originally printed in NASA Tech Briefs, Vol. 20, No. 1, January, 1996, pp. 68-69


  2. PROCEEDINGS OF THE F-8 DIGITAL FLY-BY-WIRE AND SUPERCRITICAL WING FIRST FLIGHTS' 20TH ANNIVERSARY CELEBRATION, VOLUME I
    Authors: Staff and Compiled by Kenneth E. Hodge
    Report Number: NASA-CP-3256
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A technical symposium, aircraft display dedication, and pilots' panel discussion were held on May 27, 1992, to commemorate the 20th anniversary of the first flights of the F-8 Digital Fly-By-Wire (DFBW) and Supercritical Wing (SCW) research aircraft. The symposium featured technical presentations by former key government and industry participants in the advocacy, design, aircraft modification, and flight research program activities. The DFBW and SCW technical contributions are cited. A dedication ceremony marked permanent display of both program aircraft. The panel discussion participants included eight of the eighteen research and test pilots who flew these experimental aircraft. Pilots' remarks include descriptions of their most memorable flight experiences. The report also includes a survey of the Gulf Air War, an after-dinner presentation by noted aerospace author and historian Dr. Richard Hallion.
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    Subject Category: 05
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    Report Date: February 1996
    No. Pages: 175
    Keywords:      Aerodynamics; Digital systems; F-8; Flight; Fly-by-wire; History; Research; Supercritical Wing
    Notes: See also: Volume II


  3. PROCEEDINGS OF THE F-8 DIGITAL FLY-BY-WIRE AND SUPERCRITICAL WING FIRST FLIGHTS' 20TH ANNIVERSARY CELEBRATION, VOLUME II
    Authors: Staff and Compiled by Kenneth E. Hodge
    Report Number: NASA-CP-3256
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A technical symposium, aircraft display dedication, and pilots' panel discussion were held on May 27, 1992, to commemorate the 20th anniversary of the first flights of the F-8 Digital Fly-By-Wire (DFBW) and Supercritical Wing (SCW) research aircraft. The symposium featured technical presentations by former key government and industry participants in the advocacy, design, aircraft modification, and flight research program activities. The DFBW and SCW technical contributions are cited. A dedication ceremony marked permanent display of both program aircraft. The panel discussion participants included eight of the eighteen research and test pilots who flew these experimental aircraft. Pilots' remarks include descriptions of their most memorable flight experiences. The report also includes a survey of the Gulf Air War, an after-dinner presentation by noted aerospace author and historian Dr. Richard Hallion.
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    Subject Category: 05
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    Report Date: February 1996
    No. Pages: 94
    Keywords:      Aerodynamics; Digital systems; F-8; Flight; Fly-by-wire; History; Research; Supercritical Wing
    Notes: See also: Volume I


  4. BIOMECHANICALLY INDUCED AND CONTROLLER COUPLED OSCILLATIONS EXPERIENCED ON THE F-16XL AIRCRAFT DURING ROLLING MANEUVERS
    Authors: John W. Smith and Terry Montgomery
    Report Number: NASA-TM-4752
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: During rapid rolling maneuvers, the F-16 XL aircraft exhibits a 2.5 Hz lightly damped roll oscillation, perceived and described as “roll ratcheting.” This phenomenon is common with fly-by-wire control systems, particularly when primary control is derived through a pedestal-mounted sidearm controller. Analytical studies have been conducted to model the nature of the integrated control characteristics. The analytical results complement the flight observations. A three-degree-of-freedom linearized set of aerodynamic matrices was assembled to simulate the aircraft plant. The lateral–directional control system was modeled as a linear system. A combination of two second-order transfer functions was derived to couple the lateral acceleration feedthrough effect of the operator’s arm and controller to the roll stick force input. From the combined systems, open-loop frequency responses and a time history were derived, describing and predicting an analogous in-flight situation. This report describes the primary control, aircraft angular rate, and position time responses of the F-16 XL-2 aircraft during subsonic and high-dynamic-pressure rolling maneuvers. The analytical description of the pilot’s arm and controller can be applied to other aircraft or simulations to assess roll ratcheting susceptibility.
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    Subject Category: 08
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    Report Date: July 1996
    No. Pages: 37
    Funding Organization: WU-505-68
    Keywords:      Biomechanics; F-16XL airplane; Flying qualities; Handling qualities; Human factors; Pilot induced oscillations; Roll control; Roll ratcheting


  5. A NEW CORRECTION TECHNIQUE FOR STRAIN-GAGE MEASUREMENTS ACQUIRED INTRANSIENT-TEMPERATURE ENVIRONMENTS , Technical Paper
    Authors: W. Lance Richards
    Report Number: NASA-TP-3593
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Significant strain-gage errors may exist in measurements acquired in transient-temperature environments if conventional correction methods are applied. As heating or cooling rates increase, temperature gradients between the strain-gage sensor and substrate surface increase proportionally. These temperature gradients introduce strain-measurement errors that are currently neglected in both conventional strain-correction theory and practice. Therefore, the conventional correction theory has been modified to account for these errors. A new experimental method has been developed to correct strain-gage measurements acquired in environments experiencing significant temperature transients. The new correction technique has been demonstrated through a series of tests in which strain measurements were acquired for temperature-rise rates ranging from 1 to greater than 100 degrees F/sec. Strain-gage data from these tests have been corrected with both the new and conventional methods and then compared with an analysis. Results show that, for temperature-rise rates greater than 10 degrees F/sec, the strain measurements corrected with the conventional technique produced strain errors that deviated from analysis by as much as 45 percent, whereas results corrected with the new technique were in good agreement with analytical results.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 35
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    Report Date: March 1996
    No. Pages: 34
    Funding Organization: 505-70-63
    Keywords:      Apparent strain; Measurement errors; Strain gage instrumentation; Structural testing


  6. DEVELOPMENT AND FLIGHT EVALUATION OF AN EMERGENCY DIGITAL FLIGHT CONTROL SYSTEM USING ONLY ENGINE THRUST ON AN F-15 AIRPLANE , Technical Paper
    Authors: Frank W. Burcham, Jr.,, Trindel A. Maine,, C. Gordon Fullerton and Lannie Dean Webb
    Report Number: NASA-TP-3627
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A propulsion-controlled aircraft (PCA) system for emergency flight control of aircraft with no flight controls was developed and flight tested on an F-15 airplane at the NASA Dryden Flight Research Center. The airplane has been flown in a throttles–only manual mode and with an augmented system called PCA in which pilot thumbwheel commands and aircraft feedback parameters were used to drive the throttles. Results from a 36-flight evaluation showed that the PCA system can be used to safely land an airplane that has suffered a major flight control system failure. The PCA system was used to recover from a severe upset condition, descend, and land. Guest pilots have also evaluated the PCA system. This paper describes the principles of throttles-only flight control; a history of loss-of-control accidents; a description of the F-15 airplane; the PCA system operation, simulation, and flight testing; and the pilot comments.
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    Subject Category: 08
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    Report Date: September 1996
    No. Pages: 105
    Funding Organization: WU 533-02-31
    Keywords:      Emergency flight control; F-15 airplane; Flight control systems; Propulsion-controlled aircraft; Propulsive control; Throttle-only control


  7. FLIGHT EVALUATION OF AN AIRCRAFT WITH SIDE AND CENTER STICK CONTROLLERS AND RATE-LIMITED AILERONS , Contractor Report
    Authors: P. R. Deppe, C. R. Chalk and M. F. Shafer
    Report Number: NASA-CR-198055
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: As part of an ongoing government and industry effort to study the flying qualities of aircraft with rate-limited control surface actuators, two studies were previously flown to examine an algorithm developed to reduce the tendency for pilot-induced oscillation when rate limiting occurs. This algorithm, when working properly, greatly improved the performance of the aircraft in the first study. In the second study, however, the algorithm did not initially offer as much improvement. The differences between the two studies caused concern. The study detailed in this paper was performed to determine whether the performance of the algorithm was affected by the characteristics of the cockpit controllers. Time delay and flight control system noise were also briefly evaluated. An in-flight simulator, the Calspan Learjet 25, was programmed with a low roll actuator rate limit, and the algorithm was programmed into the flight control system. Side- and center-stick controllers, force and position command signals, a rate-limited feel system, a low-frequency feel system, and a feel system damper were evaluated. The flight program consisted of four flights and 38 evaluations of test configurations. Performance of the algorithm was determined to be unaffected by using side- or center-stick controllers or force or position command signals. The rate-limited feel system performed as well as the rate-limiting algorithm but was disliked by the pilots. The low-frequency feel system and the feel system damper were ineffective. Time delay and noise were determined to degrade the performance of the algorithm.
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    Subject Category: 08
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    Report Date: November 1996
    No. Pages: 153
    Funding Organization: WU 505-64-30
    Keywords:      Flight control systems; Handling qualities; Lear 25; Pilot-induced oscillation; Rate-limiting; Variable stability
    Notes: Technical Monitor: Mary F. Shafer, NASA Dryden Flight Research Center. P. R. Deppe and C. R. Chalk, Calspan Advanced Technology Center. Previously published as Calspan Report 8091-2.


  8. STRUCTURAL DYNAMIC MODEL OBTAINED FROM FLIGHT FOR USE WITH PILOTED SIMULATION AND HANDLING QUALITIES ANALYSIS
    Authors: Bruce G. Powers
    Report Number: NASA-TM-4747
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The ability to use flight data to determine an aircraft model with structural dynamic effects suitable for piloted simulation and handling qualities analysis has been developed. This technique was demonstrated using SR71 flight test data. For the SR71 aircraft, the most significant structural response is the longitudinal first-bending mode. This mode was modeled as a second-order system, and the other higher order modes were modeled as a time delay. The distribution of the modal response at various fuselage locations was developed using a uniform beam solution, which can be calibrated using flight data. This approach was compared to the mode shape obtained from the ground vibration test, and the general form of the uniform beam solution was found to be a good representation of the mode shape in the areas of interest. To calibrate the solution, pitch-rate and normal-acceleration instrumentation is required for at least two locations. With the resulting structural model incorporated into the simulation, a good representation of the flight characteristics was provided for handling qualities analysis and piloted simulation.
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    Subject Category: 08
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    Report Date: June 1996
    No. Pages: 24
    Funding Organization: WU 537-09-22
    Keywords:      Handling qualities; Parameter estimation; Simulation; Structural dynamics
    Notes: Bruce G. Powers, Analytical Services and Material, Inc., Edwards, California


  9. A HISTORICAL PERSPECTIVE OF THE YF-12A THERMAL LOADS AND STRUCTURES PROGRAM
    Authors: Jerald M. Jenkins and Robert D. Quinn
    Report Number: NASA-TM-104317
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Around 1970, the YF-12A loads and structures efforts focused on numerous technological issues that needed defining with regard to aircraft that incorporate hot structures in the design. Laboratory structural heating test technology with infrared systems was largely created during this program. The program demonstrated the ability to duplicate the complex flight temperatures of an advanced super-sonic airplane in a ground-based laboratory. The ability to heat and load an advanced operational air-craft in a laboratory at high temperatures and return it to flight status without adverse effects was dem-onstrated. The technology associated with measuring loads with strain gages on a hot structure was demonstrated with a thermal calibration concept. The results demonstrated that the thermal stresses were significant although the airplane was designed to reduce thermal stresses. Considerable model-ing detail was required to predict the heat transfer and the corresponding structural characteristics. The overall YF-12A research effort was particularly productive, and a great deal of flight, laboratory, test and computational data were produced and cross-correlated.
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    Subject Category: 01
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    Report Date: May 1996
    No. Pages: 34
    Funding Organization: WU 505-70-63
    Keywords:      F-111 aircraft; Flight loads; Heat transfer; Hot structures; Loads measurement; Strain gages; Thermal loads and structures; YF-12A aircraft 35


  10. AEROSERVOELASTIC MODELING AND VALIDATION OF A THRUST-VECTORING F/A-18 AIRCRAFT , Technical Paper
    Authors: Martin J. Brenner
    Report Number: NASA-TP-3647
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An F/A-18 aircraft was modified to perform flight research at high angles of attack (AOA) using thrust vectoring and advanced control law concepts for agility and performance enhancement and to provide a testbed for the computational fluid dynamics community. Aeroservoelastic (ASE) characteristics had changed considerably from the baseline F/A18 aircraft because of structural and flight control system amendments, so analyses and flight tests were performed to verify structural stability at high AOA. Detailed actuator models that consider the physical, electrical, and mechanical elements of actuation and its installation on the airframe were employed in the analysis to accurately model the coupled dynamics of the airframe, actuators, and control surfaces. This report describes the ASE modeling procedure, ground test validation, flight test clearance, and test data analysis for the reconfigured F/A18 aircraft. Multivariable ASE stability margins are calculated from flight data and compared to analytical margins. Because this thrust-vectoring configuration uses exhaust vanes to vector the thrust, the modeling issues are nearly identical for modern multiaxis nozzle configurations. This report correlates analysis results with flight test data and makes observations concerning the application of the linear predictions to thrust-vectoring and high-AOA flight.
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    Subject Category: 08
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    Report Date: September 1996
    No. Pages: 64
    Funding Organization: 505-68-30
    Keywords:      Aeroservoelasticity; Aircraft vibration; Dynamic stability; F/A-18 aircraft; Flight dynamics; Structural dynamics; Thrust vectoring


  11. A DYNAMIC RESPONSE MODEL FOR PRESSURE SENSORS IN CONTINUUM AND HIGH KNUDSEN NUMBER FLOWS WITH LARGE TEMPERATURE GRADIENTS , Technical Memorandum
    Authors: Stephen A. Whitmore, Brian J. Petersen and David D. Scott
    Report Number: NASA-TM-4728
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper develops a dynamic model for pressure sensors in continuum and rarefied flows with longitudinal temperature gradients. The model was developed from the unsteady Navier-Stokes momentum, energy, and continuity equations and was linearized using small perturbations. The energy equation was decoupled from momentum and continuity assuming a polytropic flow process. Rarefied flow conditions were accounted for using a slip flow boundary condition at the tubing wall. The equations were radially averaged and solved assuming gas properties remain constant along a small tubing element. This fundamental solution was used as a building block for arbitrary geometries where fluid properties may also vary longitudinally in the tube. The problem was solved recursively starting at the transducer and working upstream in the tube. Dynamic frequency response tests were performed for continuum flow conditions in the presence of temperature gradients. These tests validated the recursive formulation of the model. Model steady-state behavior was analyzed using the final value theorem. Tests were performed for rarefied flow conditions and compared to the model steady-state response to evaluate the regime of applicability. Model comparisons were excellent for Knudsen numbers up to 0.6. Beyond this point, molecular affects caused model analyses to become inaccurate.
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    Subject Category: 02
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    Report Date: January 1996
    No. Pages: 42
    Funding Organization: WU 52-00-RR-00-000
    Keywords:      Hypersonic aerodynamics; Knudsen number; Pneumatic attenuation; Pressure sensing;
    Notes: Presented as AIAA 96-0563 at the 34th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada,Jan. 15–18, 1996. Stephen A. Whitmore, Dryden Flight Research Center, Edwards, California; Brian J. Petersen, UCLA,Los Angeles, California; David D. Scott, L


  12. DEVELOPMENT OF A REAL-TIME TRANSPORT PERFORMANCE OPTIMIZATION METHODOLOGY , Technical Memorandum
    Authors: Glenn Gilyard
    Report Number: NASA-TM-4730
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The practical application of real-time performance optimization is addressed (using a wide-body transport simulation) based on real-time measurements and calculation of incremental drag from forced response maneuvers. Various controller combinations can be envisioned although this study used symmetric outboard aileron and stabilizer. The approach is based on navigation instrumentation and other measurements found on state-of-the-art transports. This information is used to calculate winds and angle of attack. Thrust is estimated from a representative engine model as a function of measured variables. The lift and drag equations are then used to calculate lift and drag coefficients. An expression for drag coefficient, which is a function of parasite drag, induced drag, and aileron drag, is solved from forced excitation response data. Estimates of the parasite drag, curvature of the aileron drag variation, and minimum drag aileron position are produced. Minimum drag is then obtained by repositioning the symmetric aileron. Simulation results are also presented which evaluate the affects of measurement bias and resolution.
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    Subject Category: 01, 02 and 03
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    Report Date: January 1996
    No. Pages: 18
    Funding Organization: WU 505-64-1024
    Keywords:      Aircraft performance; Airfoil and wing aerodynamics; Cambered wings; Drag reduction; Fuel consumption; Flight optimization; F-111 aircraft; Optimal control; Optimization
    Notes: Presented as AIAA 96-0093 at the 34th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 15–18, 1996.


  13. ANALYTICAL AND EXPERIMENTAL VERIFICATION OF A FLIGHT ARTICLE FOR A MACH-8 BOUNDARY-LAYER EXPERIMENT , Technical Memorandum
    Authors: W. Lance Richards and Richard C. Monaghan
    Report Number: NASA-TM-4733
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Preparations for a boundary-layer transition experiment to be conducted on a future flight mission of the air-launched Pegasus® rocket are underway. The experiment requires a flight-test article called a glove to be attached to the wing of the Mach-8 first-stage booster. A three-dimensional, nonlinear finite-element analysis has been performed and significant small-scale laboratory testing has been accomplished to ensure the glove design integrity and quality of the experiment. Reliance on both the analysis and experiment activities has been instrumental in the success of the flight-article design. Results obtained from the structural analysis and laboratory testing show that all glove components are well within the allowable thermal stress and deformation requirements to satisfy the experiment objectives.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 39
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    Report Date: March 1996
    No. Pages: 15
    Funding Organization: WU 505-70-59
    Keywords:      Computational methods; Flight test techniques; Structural analysis; Structural design; Structural testing
    Notes: Presented at the First International Conference on Computational Methods and Testing for Engineering Integrity, Kuala Lumpur, Malaysia, March 19–21, 1996.


  14. THE X-31A QUASI-TAILLESS FLIGHT TEST RESULTS , Technical Paper
    Authors: John T. Bosworth and P. C. Stoliker
    Report Number: NASA-TP-3624
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A quasi-tailless flight investigation was launched using the X-31A enhanced fighter maneuverability airplane. In-flight simulations were used to assess the effect of partial to total vertical tail removal. The rudder control surface was used to cancel the stabilizing effects of the vertical tail, and yaw thrust vector commands were used to restabilize and control the airplane. The quasi-tailless mode was flown supersonically with gentle maneuvering and subsonically in precision approaches and ground attack profiles. Pilot ratings and a full set of flight test measurements were recorded. This report describes the results obtained and emphasizes the lessons learned from the X-31A flight test experiment. Sensor-related issues and their importance to a quasi-tailless simulation and to ultimately controlling a directionally unstable vehicle are assessed. The X-31A quasi-tailless flight test experiment showed that tailless and reduced tail fighter aircraft are definitely feasible. When the capability is designed into the airplane from the beginning, the benefits have the potential to outweigh the added complexity required.
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    Subject Category: 08
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    Report Date: June 1996
    No. Pages: 42
    Funding Organization: WU-505-68-30
    Keywords:      Aircraft stability; Flight test; Flight research; Flight simulation; Quasi-taillessairplanes; Rudder control; Tailless aircraft; Thrust vector; X-31 Airplane; Yaw control
    Notes: Presented at the Second Test and Evaluation International Aerospace Forum, London, England,June 25–27, 1996.


  15. LONGITUDINAL EMERGENCY CONTROL SYSTEM USING THRUST MODULATION DEMONSTRATION ON AN MD-11 AIRPLANE , Conference Paper
    Authors: John J. Burken, Trindel A. Maine, Frank W. Burcham, Jr. and Jeffery A. Kahler
    Report Number: AIAA-96-3062
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This report describes how an MD-11 airplane landed using only thrust modulation, with the control surfaces locked. The propulsion-controlled aircraft system would be used if the aircraft suffered a major primary flight control system failure and lost most or all of the hydraulics. The longitudinal and lateral-directional controllers were designed and flight tested, but only the longitudinal control of flightpath angle is addressed in this paper. A flight-test program was conducted to evaluate the aircraft's high-altitude flying characteristics and to demonstrate its capacity to perform safe landings. In addition, over 50 low approaches and three landings without the movement of any aerodynamic control surfaces were performed. The longitudinal control modes include a wing engines only mode for flightpath control and a three-engine operation mode with speed control and dynamic control of the flightpath angle using the tail engine. These modes were flown in either a pilot-commanded mode or an instrument landing system coupled mode. Also included are the results of an analytical study of an autothrottle longitudinal controller designed to improve the phugoid damping. This mode requires the pilot to use differential throttles for lateral control.
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    Subject Category: 08
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    Report Date: July 1996
    No. Pages: 17
    Funding Organization: WU 505-64-10
    Keywords:      Control system failure; Hydraulic system failure; Longitudinal emergency controller using thrust; MD-11 airplane; Propulsion control
    Notes: Presented at the 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Lake Buena Vista, Florida, July 1–3, 1996, J.J. Burken, T.A. Maine, F.W. Burcham, Jr., NASA DFRC, Edwards, CA; J.A. Kahler, Honeywell, Inc., Phoenix, AZ


  16. THERMAL AND MECHANICAL BUCKLING ANALYSIS OF HYPERSONIC AIRCRAFT HAT-STIFFENED PANELS WITH VARYING FACE SHEET GEOMETRY AND FIBER ORIENTATION , Technical Memorandum
    Authors: William L. Ko
    Report Number: NASA-TM-4770
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Mechanical and thermal buckling behavior of monolithic and metal-matrix composite hat-stiffened panels were investigated. The panels have three types of face-sheet geometry: flat face sheet, microdented face sheet, and microbulged face sheet. The metal-matrix composite panels have three types of face-sheet layups, each of which is combined with various types of hat composite layups. Finite-element method was used in the eigenvalue extractions for both mechanical and thermal buckling. The thermal buckling analysis required both eigenvalue and material property iterations. Graphical methods of the dual iterations are shown. The mechanical and thermal buckling strengths of the hat-stiffened panels with different face-sheet geometry are compared. It was found that by just microdenting or microbulging of the face sheet, the axial, shear, and thermal buckling strengths of both types of hat-stiffened panels could be enhanced considerably. This effect is more conspicuous for the monolithic panels. For the metal-matrix composite panels, the effect of fiber orientations on the panel buckling strengths was investigated in great detail, and various composite layup combinations offering high panel buckling strengths are presented. The axial buckling strength of the metal-matrix panel was sensitive to the change of hat fiber orientation. However, the lateral, shear, and thermal buckling strengths were insensitive to the change of hat fiber orientation.
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    Subject Category: 39
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    Report Date: December 1996
    No. Pages: 36
    Funding Organization: WU 505-63-50
    Keywords:      Hat-stiffened panels; Mechanical buckling; Metal-matrix composites; Thermal buckling; Varying face sheet geometry; Varying fiber orientation
    Notes: n.a.


  17. A FORMAL ALGORITHM FOR ROUTING TRACES ON A PRINTED CIRCUIT BOARD , Technical Paper
    Authors: David R. Hedgley, Jr.
    Report Number: NASA-TP-3639
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper addresses the classical problem of printed circuit board routing: that is, the problem of automatic routing by a computer other than by brute force that causes the execution time to grow exponentially as a function of the complexity. Most of the present solutions are either inexpensive but not efficient and fast, or efficient and fast but very costly. Many solutions are proprietary, so not much is written or known about the actual algorithms upon which these solutions are based. This paper presents a formal algorithm for routing traces on a printed circuit board. The solution presented is very fast and efficient and for the first time speaks to the question eloquently by way of symbolic statements.
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    Subject Category: 61
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    Report Date: September 1996
    No. Pages: 12
    Funding Organization: WU-505-68-84
    Keywords:      Artificial intelligence; Printed circuit board; Routing
    Notes: n.a.


  18. DEVELOPMENT AND FLIGHT TEST OF AN AUGMENTED THRUST-ONLY FLIGHT CONTROL SYSTEM ON AN MD11 TRANSPORT AIRPLANE
    Authors: Frank W. Burcham, Jr., Trindel A. Maine, John J. Burken and Drew Pappas
    Report Number: NASA-TM-4745
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An emergency flight control system using only engine thrust, called Propulsion-Controlled Aircraft (PCA), has been developed and flight tested on an MD11 airplane. In this thrust-only control system, pilot flightpath and track commands and aircraft feedback parameters are used to control the throttles. The PCA system was installed on the MD11 airplane using software modifications to existing computers. Flight test results 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 and altitudes. The PCA approaches, go-arounds, and three landings without the use of any normal flight controls have been demonstrated, including instrument landing system-coupled hands-off landings. The PCA operation was used to recover from an upset condition. In addition, PCA was tested at altitude with all three hydraulic systems turned off. This paper reviews the principles of throttles-only flight control; describes the MD11 airplane and systems; and discusses PCA system development, operation, flight testing, and pilot comments.
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    Subject Category: 08
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          Postscript (2,935 KBytes)
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    Report Date: July 1996
    No. Pages: 28
    Funding Organization: WU 505-68-10
    Keywords:      Emergency flight control; MD11 airplane; Propulsion-controlled aircraft; Propulsive control; Throttle-only control
    Notes: Presented as AIAA 963742, AIAA Guidance, Navigation, and Control Conference, San Diego, California, July 29–31, 1996. F.W. Burcham, Jr., T.A. Maine, and J.J. Burken, NASA Dryden Flight Research Center, Edwards. California. Pappas, McDonnell Douglas Aerosp


  19. WIND-TUNNEL DEVELOPMENT OF AN SR-71 AEROSPIKE ROCKET FLIGHT TEST CONFIGURATION , Technical Memorandum
    Authors: Timothy R. Moes, Brent R. Cobleigh, Timothy R. Connors, Timothy H. Cox, Stephen C. Smith and Norm Shirakata
    Report Number: NASA-TM-4749
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight experiment has been proposed to investigate the performance of an aerospike rocket motor installed in a lifting body configuration. An SR-71 airplane would be used to carry the aerospike configuration to the desired flight test conditions. Wind-tunnel tests were completed on a 4-percent scale SR-71 airplane with the aerospike pod mounted in various locations on the upper fuselage. Testing was accomplished using sting and blade mounts from Mach 0.6 to Mach 3.2. Initial test objectives included assessing transonic drag and supersonic lateral–directional stability and control. During these tests, flight simulations were run with wind-tunnel data to assess the acceptability of the configurations. Early testing demonstrated that the initial configuration with the aerospike pod near the SR-71 center of gravity was unsuitable because of large nosedown pitching moments at transonic speeds. The excessive trim drag resulting from accommodating this pitching moment far exceeded the excess thrust capability of the airplane. Wind-tunnel testing continued in an attempt to find a configuration suitable for flight test. Multiple configurations were tested. Results indicate that an aft-mounted model configuration possessed acceptable performance, stability, and control characteristics.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (632 KBytes)
          PDF (364 KBytes)
    Report Date: January 1996
    No. Pages: 26
    Funding Organization: WU 505-68-52
    Keywords:      Aerodynamic models; Aerospike; Flight simulations; Rocket; SR-71 airplane; Wind tunnel
    Notes: Presented as AIAA 96-2409 at the 14th Applied Aerodynamics Conf., LA, June 1996. T. Moes, B. Cobleigh, T. Conners, T. Cox, NASA Dryden Flight Research Ctr, CA; Stephen Smith, NASA Ames Research Ctr, CA; Norm Shirakata, Lockheed-Martin Skunk Works, CA


  20. LONGITUDINAL EMERGENCY CONTROL SYSTEM USING THRUST MODULATION DEMONSTRATION ON AN MD-11 AIRPLANE , Technical Memorandum
    Authors: John J. Burken, Trindel A. Maine, Frank W. Burcham, Jr. and Jeffery A. Kahler
    Report Number: NASA-TM-104318
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This report describes how an MD-11 airplane landed using only thrust modulation, with the control surfaces locked. The propulsion-controlled aircraft system would be used if the aircraft suffered a major primary flight control system failure and lost most or all of the hydraulics. The longitudinal and lateral-directional controllers were designed and flight tested, but only the longitudinal control of flightpath angle is addressed in this paper. A flight-test program was conducted to evaluate the aircraft's high-altitude flying characteristics and to demonstrate its capacity to perform safe landings. In addition, over 50 low approaches and three landings without the movement of any aerodynamic control surfaces were performed. The longitudinal control modes include a wing engines only mode for flightpath control and a three-engine operation mode with speed control and dynamic control of the flightpath angle using the tail engine. These modes were flown in either a pilot-commanded mode or an instrument landing system coupled mode. Also included are the results of an analytical study of an autothrottle longitudinal controller designed to improve the phugoid damping. This mode requires the pilot to use differential throttles for lateral control.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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          Postscript (523 KBytes)
          PDF (254 KBytes)
    Report Date: July 1996
    No. Pages: 17
    Funding Organization: WU 505-64-10
    Keywords:      Control system failure; Hydraulic system failure; Longitudinal emergency controller using thrust; MD-11 airplane; Propulsion control
    Notes: Presented as AIAA 96-3062 at the 32nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Lake Buena Vista, Florida, July 1–3, 1996, J.J. Burken, T.A. Maine, F.W. Burcham, Jr., NASA DFRC, Edwards, CA; J.A. Kahler, Honeywell, Inc., Phoenix, AZ


  21. FLIGHT TEST OF A PROPULSION-BASED EMERGENCY CONTROL SYSTEM ON THE MD-11 AIRPLANE WITH EMPHASIS ON THE LATERAL AXIS , Technical Memorandum
    Authors: John J. Burken, Frank W. Burcham, Jr., Trindel A. Maine, John Feather, Steven Goldthorpe and Jeffrey A. Kahler
    Report Number: NASA-TM-4746
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A large, civilian, multiengine transport MD-11 airplane control system was recently modified to perform as an emergency backup controller using engine thrust only. The emergency backup system, referred to as the propulsion-controlled aircraft (PCA) system, would be used if a major primary flight control system fails. To allow for longitudinal and lateral–directional control, the PCA system requires at least two engines and is implemented through software modifications. A flight-test program was conducted to evaluate the PCA system high-altitude flying characteristics and to demonstrate its capacity to perform safe landings. The cruise flight conditions, several low approaches and one landing without any aerodynamic flight control surface movement, were demonstrated. This paper presents results that show satisfactory performance of the PCA system in the longitudinal axis. Test results indicate that the lateral–directional axis of the system performed well at high altitude but was sluggish and prone to thermal upsets during landing approaches. Flight-test experiences and test techniques are also discussed with emphasis on the lateral–directional axis because of the difficulties encountered in flight test.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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          Postscript (683 KBytes)
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    Report Date: July 1996
    No. Pages: 15
    Funding Organization: WU 505-64-10
    Keywords:      Control system failure; Emergency controller using thrust; Hydraulic system failure; MD-11 airplane; Propulsion-only control
    Notes: Presented as AIAA 96-3919 at the AIAA Guidance, Navigation, and Control Conference, San Diego, California, July 29–31, 1996.


  22. DESIGN AND LABORATORY VALIDATION OF A CAPACITIVE SENSOR FOR MEASURING THE RECESSION OF A THIN-LAYERED ABLATOR
    Authors: Gregory K. Noffz and Michael P. Bowman
    Report Number: NASA-TM-4777
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight vehicles are typically instrumented with subsurface thermocouples to estimate heat transfer at the surface using inverse analysis procedures. If the vehicle has an ablating heat shield, however, temperature time histories from subsurface thermocouples no longer provide enough information to estimate heat flux at the surface. In this situation, the geometry changes and thermal energy leaves the surface in the form of ablation products. The ablation rate is required to estimate heat transfer to the surface. A new concept for a capacitive sensor has been developed to measure ablator depth using the ablator’s dielectric effect on a capacitor’s fringe region. Relying on the capacitor’s fringe region enables the gage to be flush mounted in the vehicle’s permanent structure and not intrude into the ablative heat shield applied over the gage. This sensor’s design allows nonintrusive measurement of the thickness of dielectric materials, in particular, the recession rates of low-temperature ablators applied in thin (0.020 to 0.060 in. (0.05 to 0.15 mm)) layers. Twenty capacitive gages with 13 different sensing element geometries were designed, fabricated, and tested. A two-dimensional finite-element analysis was performed on several candidate geometries. Calibration procedures using ablator-simulating shims are described. A one-to-one correspondence between system output and dielectric material thickness was observed out to a thickness of 0.055 in. (1.4 mm) for a material with a permittivity about three times that of air or vacuum. A novel method of monitoring the change in sensor capacitance was developed. This technical memorandum suggests further improvements in gage design and fabrication techniques.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 34
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    Report Date: November 1996
    No. Pages: 38
    Funding Organization: WU 505-68-30
    Keywords:      Ablation; Ablation measurement; Capacitive sensors; Fringe field; Thin film measurement


  23. F-15B/FLIGHT TEST FIXTURE II: A TEST BED FOR FLIGHT RESEARCH , Technical Memorandum
    Authors: David. M Richwine
    Report Number: NASA-TM-4782
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden Flight Research Center has developed a second-generation flight test fixture for use as a generic test bed for aerodynamic and fluid mechanics research. The Flight Test Fixture II (FTF–II) is a low-aspect-ratio vertical fin-like shape that is mounted on the centerline of the F-15B lower fuselage. The fixture is designed for flight research at Mach numbers to a maximum of 2.0. The FTF–II is a composite structure with a modular configuration and removable components for functional flexibility. This report documents the flow environment of the fixture, such as surface pressure distributions and boundary-layer profiles, throughout a matrix of conditions within the F-15B/FTF–II flight envelope. Environmental conditions within the fixture are presented to assist in the design and testing of future avionics and instrumentation. The intent of this document is to serve as a user’s guide and assist in the development of future flight experiments that use the FTF–II as a test bed. Additional information enclosed in the appendices has been included to assist with more detailed analyses, if required.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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          Postscript (2,030 KBytes)
          PDF (1,119 KBytes)
    Report Date: December 1996
    No. Pages: 52
    Funding Organization: WU 505-68-52
    Keywords:      Aerodynamic characterization; Boundary-layer pressures; F-15B aerodynamic test bed; F15 directional stability; Flight test facility; Flight test fixture; Flow visualization; Surface pressure distributions


  24. FUTURE FLIGHT TEST PLANS OF AN AXISYMMETRIC HYDROGEN-FUELED SCRAMJET ENGINE ON THE HYPERSONIC FLYING LABORATORY , AIAA-96-4572
    Authors: Alexander S. Roudakov, Vyacheslav L. Semenov, Valeriy I. Kopchenov and John W. Hicks
    Report Number: H-2115
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Under a contract with NASA, a joint Central Institute of Aviation Motors (CIAM) and NASA team is preparing to conduct the fourth flight test of a dual-mode scramjet aboard the CIAM Hypersonic Flying Laboratory, “Kholod.” Ground-launch, rocket boosted by a modified Russian SA-5 missile, the redesigned scramjet is to be accelerated to a new maximum velocity of Mach 6.5. This should allow for the first-time measurement of the fully supersonic combustion mode. The primary program objective is the flight-to-ground correlation of measured data with preflight analysis and wind-tunnel tests in Russia and potentially in the United States. This paper describes the development and objectives of the program as well as the technical details of the scramjet and SA-5 redesign to achieve the Mach 6.5 aim test condition. The purpose and value of a joint Russian-American program to attain overall hypersonic air-breathing technology objectives are discussed. Finally, the current project status and schedules to reach the final flight launch are discussed.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
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          Postscript (682 KBytes)
          PDF (332 KBytes)
    Report Date: November 1996
    No. Pages: 9
    Funding Organization: n.a.
    Keywords:      n.a.
    Notes: Prepared for AIAA 7th International Spaceplanes and Hypersonics Systems & Technology Conference, Nov. 18–22, 1996, VA. A.S. Roudakov, V.L. Semenov, V.I. Kopchenov, Central Institute of Aviation Motors, Moscow, Russia. J.W. Hicks, NASA DFRC, Edwards, CA.


  25. COHERENT LIDAR TURBULENCE MEASUREMENT FOR GUST LOAD ALLEVIATION , Technical Memorandum
    Authors: David Soreide, Rodney K. Bogue, L.J. Ehernberger and Hal Bagley
    Report Number: NASA-TM-104318
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Atmospheric turbulence adversely affects operation of commercial and military aircraft and is a design constraint. The airplane structure must be designed to survive the loads imposed by turbulence. Reducing these loads allows the airplane structure to be lighter, a substantial advantage for a commercial airplane. Gust alleviation systems based on accelerometers mounted in the airplane can reduce the maximum gust loads by a small fraction. These systems still represent an economic advantage. The ability to reduce the gust load increases tremendously if the turbulent gust can be measured before the airplane encounters it. A lidar system can make measurements of turbulent gusts ahead of the airplane, and the NASA Airborne Coherent Lidar for Advanced In-Flight Measurements (ACLAIM) program is developing such a lidar. The ACLAIM program is intended to develop a prototype lidar system for use in feasibility testing of gust load alleviation systems and other airborne lidar applications, to define applications of lidar with the potential for improving airplane performance, and to determine the feasibility and benefits of these applications. This paper gives an overview of the ACLAIM program, describes the lidar architecture for a gust alleviation system, and describes the prototype ACLAIM lidar system.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
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          Postscript (240 KBytes)
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    Report Date: January 1996
    No. Pages: 17
    Funding Organization: WU 505-69-59
    Keywords:      Aircraft operation; Aircraft sensors; Backscatter; Gust; Lidar; Load alleviation;Turbulence
    Notes: Presented as SPIE paper 2832-05 at the SPIE 1996 International Symposium on Optical Science, Engineering, and Instrumentation, August 4–9, 1996, Denver, Colorado. D. Soreide, Technical Consultant, Seattle, Washington. R.K. Bogue and L.J. Ehernberger, NASA


  26. X-29A LATERAL–DIRECTIONAL STABILITY AND CONTROL DERIVATIVES EXTRACTED FROM HIGH-ANGLE-OF-ATTACK FLIGHT DATA , Technical Paper
    Authors: Kenneth W. Iliff and Kon-Sheng Charles Wang
    Report Number: NASA-TP-3664
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The lateral–directional stability and control derivatives of the X-29A number 2 are extracted from flight data over an angle-of-attack range of 4 deg to 53 deg using a parameter identification algorithm. The algorithm uses the linearized aircraft equations of motion and a maximum likelihood estimator in the presence of state and measurement noise. State noise is used to model the uncommanded forcing function caused by unsteady aerodynamics over the aircraft at angles of attack above 15 deg. The results supported the flight-envelope-expansion phase of the X-29A number 2 by helping to update the aerodynamic mathematical model, to improve the real-time simulator, and to revise flight control system laws. Effects of the aircraft high gain flight control system on maneuver quality and the estimated derivatives are also discussed. The derivatives are plotted as functions of angle of attack and compared with the predicted aerodynamic database. Agreement between predicted and flight values is quite good for some derivatives such as the lateral force due to sideslip, the lateral force due to rudder deflection, and the rolling moment due to roll rate. The results also show significant differences in several important derivatives such as the rolling moment due to sideslip, the yawing moment due to sideslip, the yawing moment due to aileron deflection, and the yawing moment due to rudder deflection.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 08
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          PDF (584 KBytes)
    Report Date: December 1996
    No. Pages: 39
    Funding Organization: WU 505 68 50 00R
    Keywords:      Aerodynamic characteristics; Forward-swept wing; High angle of attack; Lateral-directional; Maximum likelihood estimation; Parameter identification; Stability and control derivatives; X-29A


  27. FLIGHT AND STATIC EXHAUST FLOW PROPERTIES OF AN F110-GE-129 ENGINE IN AN F-16XL AIRPLANE DURING ACOUSTIC TESTS , Technical Memorandum
    Authors: Jon K. Holzman, Lannie D. Webb and Frank W. Burcham, Jr.
    Report Number: NASA-TM-104326
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The exhaust flow properties (mass flow, pressure, temperature, velocity, and Mach number) of the F110-GE-129 engine in an F-16XL airplane were determined from a series of flight tests flown at NASA Dryden Flight Research Center, Edwards, California. These tests were performed in conjunction with NASA Langley Research Center, Hampton, Virginia (LaRC) as part of a study to investigate the acoustic characteristics of jet engines operating at high nozzle pressure conditions. The range of interest for both objectives was from Mach 0.3 to Mach 0.9. NASA Dryden flew the airplane and acquired and analyzed the engine data to determine the exhaust characteristics. NASA Langley collected the flyover acoustic measurements and correlated these results with their current predictive codes. This paper describes the airplane, tests, and methods used to determine the exhaust flow properties and presents the exhaust flow properties. No acoustics results are presented.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 07
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          Postscript (581 KBytes)
          PDF (349 KBytes)
    Report Date: November 1996
    No. Pages: 31
    Funding Organization: WU 505-68-10
    Keywords:      Acoustics; F110 engine; F-16XL airplane; Flyover noise
    Notes: n.a.


  28. AN OVERVIEW OF CONTROLS AND FLYING QUALITIES TECHNOLOGY ON THE F/A-18 HIGH ALPHA RESEARCH VEHICLE , Conference Paper
    Authors: Joseph W. Pahle, Keith D. Wichman, John V. Foster and W. Thomas Bundick
    Report Number: NASA-H-2123
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA F/A-18 High Alpha Research Vehicle (HARV) has been the flight test bed of a focused technology effort to significantly increase maneuvering capability at high angles of attack. Development and flight test of control law design methodologies, handling qualities metrics, performance guidelines, and flight evaluation maneuvers are described. The HARV has been modified to include two research control effectors, thrust vectoring, and actuated forebody strakes in order to provide increased control power at high angles of attack. A research flight control system has been used to provide a flexible, easily modified capability for high-angle-of-attack research controls. Different control law design techniques have been implemented and flight-tested, including eigenstructure assignment, variable gain output feedback, pseudo controls, and model-following. Extensive piloted simulation has been used to develop nonlinear performance guidelines and handling qualities criteria for high angles of attack. This paper reviews the development and evaluation of technologies useful for high-angle-of-attack control. Design, development, and flight test of the research flight control system, control laws, flying qualities specifications, and flight test maneuvers are described. Flight test results are used to illustrate some of the lessons learned during flight test and handling qualities evaluations.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: n.a.
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    Report Date: September 1996
    No. Pages: 27
    Funding Organization: n.a.
    Keywords:      n.a.
    Notes: Presented at the 5th High-Angle-of-Attack Technology Conference, NASA Langley Research Center, VA, Sept. 1996. J.W. Pahle and K.D. Wichman, NASA Dryden, CA. J.V. Foster and W.T. Bundick, NASA Langley, VA.


  29. ESTIMATING ENGINE AIRFLOW IN GAS-TURBINE POWERED AIRCRAFT WITH CLEAN AND DISTORTED INLET FLOWS , Contractor Report
    Authors: J. G. Williams, W. G. Steenken and A. J. Yuhas
    Report Number: NASA-CR-198052
    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 impact of inlet-generated total-pressure distortion on estimating levels of engine airflow. Five airflow estimation methods were studied. The Reference Method was a fan corrected airflow to fan corrected speed calibration from an uninstalled engine test. In-flight airflow estimation methods utilized the average, or individual, inlet duct static- to total-pressure ratios, and the average fan-discharge static-pressure to average inlet total-pressure ratio. Correlations were established at low distortion conditions for each method relative to the Reference Method. A range of distorted inlet flow conditions were obtained from –10 deg to +60 deg angle of attack and –7 deg to +11 deg angle of sideslip. The individual inlet duct pressure ratio correlation resulted in a 2.3 percent airflow spread for all distorted flow levels with a bias error of –0.7 percent. The fan discharge pressure ratio correlation gave results with a 0.6 percent airflow spread with essentially no systematic error. Inlet-generated total-pressure distortion and turbulence had no significant impact on the F404-GE-400 engine airflow pumping. Therefore, a speed-flow relationship may provide the best airflow estimate for a specific engine under all flight conditions.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 07
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          Postscript (2,815 KBytes)
          PDF (5,489 KBytes)
    Report Date: September 1996
    No. Pages: 79
    Funding Organization: WU 505-68-30
    Keywords:      Airflow correlation techniques; Airflow errors; Distorted inlet airflow; Engine airflow; GE 404 engine airflow
    Notes: DFRC Tech Monitor: K. Walsh. J.G. Williams & W.G. Steenken, GE Aircraft Engines, Cincinnati, OH. A.J. Yuhas, AS&M, Inc., Hampton, VA. Presented at the High-Angle-of-Attack Tech Conf, NASA Langley, Hampton, VA, Sept. 17–19, 1996.


  30. EVALUATION OF HIGH-ANGLE-OF-ATTACK HANDLING QUALITIES FOR THE X-31A USING STANDARD EVALUATION MANEUVERS
    Authors: Patrick C. Stoliker and John T. Bosworth
    Report Number: NASA-TM-104322
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The X-31A aircraft gross-acquisition and fine-tracking handling qualities have been evaluated using standard evaluation maneuvers developed by Wright Laboratory, Wright-Patterson Air Force Base. The emphasis of the testing is in the angle-of-attack range between 30 and 70 degrees. Longitudinal gross-acquisition handling qualities results show borderline Level 1/Level 2 performance. Lateral gross-acquisition testing results in Level 1/Level 2 ratings below 45 degrees angle of attack, degrading into Level 3 as angle of attack increases. The fine-tracking performance in both longitudinal and lateral axes also receives Level 1 ratings near 30 degrees angle of attack, with the ratings tending towards Level 3 at angles of attack greater than 50 degrees. These ratings do not match the expectations from the extensive close-in combat testing where the X-31A aircraft demonstrated fair to good handling qualities maneuvering for high angles of attack. This paper presents the results of the high-angle-of-attack handling qualities flight testing of the X-31A aircraft. Discussion of the preparation for the maneuvers, the pilot ratings, and selected pilot comments are included. Evaluation of the results is made in conjunction with existing Neal-Smith, bandwidth, Smith-Geddes, and military specifications.
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    Subject Category: 08
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    Report Date: September 1996
    No. Pages: 36
    Funding Organization: WU 505-68-30
    Keywords:      Flight controls; Handling qualities; High angle of attack; Pilot-induced oscillation; Standard evaluation maneuvers; X-31A aircraft
    Notes: Presented at the High-Angle-of-Attack Technology Conference, NASA Langley Research Center, Hampton, Virginia, Sept. 17–19, 1996. Also presented at the AGARD Flight Vehicle Integration Panel Symposium on Advances in Flight Testing, Lisbon, Portugal, Sept.


  31. DYNAMIC INLET DISTORTION PREDICTION WITH A COMBINED COMPUTATIONAL FLUID DYNAMICS AND DISTORTION SYNTHESIS APPROACH , Contractor Report
    Authors: W. P. Norby, J. A. Ladd and A. J. Yuhas
    Report Number: NASA-CR-198053
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A procedure has been developed for predicting peak dynamic inlet distortion. This procedure combines Computational Fluid Dynamics (CFD) and distortion synthesis analysis to obtain a prediction of peak dynamic distortion intensity and the associated instantaneous total pressure pattern. A prediction of the steady state total pressure pattern at the Aerodynamic Interface Plane is first obtained using an appropriate CFD flow solver. A corresponding inlet turbulence pattern is obtained from the CFD solution via a correlation linking root mean square (RMS) inlet turbulence to a formulation of several CFD parameters representative of flow turbulence intensity. This correlation was derived using flight data obtained from the NASA High Alpha Research Vehicle flight test program and several CFD solutions at conditions matching the flight test data. A distortion synthesis analysis is then performed on the predicted steady state total pressure and RMS turbulence patterns to yield a predicted value of dynamic distortion intensity and the associated instantaneous total pressure pattern.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 07
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    Report Date: September 1996
    No. Pages: 38
    Funding Organization: WU 505-69-30
    Keywords:      Computational fluid dynamics; Distortion synthesis; F/A-18 airplane; Inlet; Inlet distortion
    Notes: Dryden Technical Monitor: Kevin Walsh. W.P. Norby and J.A. Ladd, McDonnell Douglas Aerospace, St. Louis, MO, and A.J. Yuhas, AS&M, Inc., Hampton, VA. Presented at the High-Angle-of-Attack Tech Conf, NASA Langley, Hampton, VA, Sept. 17–19, 1996.


  32. TITANIUM HONEYCOMB PANEL TESTING , Technical Memorandum
    Authors: W. Lance Richards and Randolph C. Thompson
    Report Number: NASA-TM-4768
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Thermal–mechanical tests were performed on a titanium honeycomb sandwich panel to experimentally validate the hypersonic wing panel concept and compare test data with analysis. Details of the test article, test fixture development, instrumentation, and test results are presented. After extensive testing to 900 deg F, non-destructive evaluation of the panel has not detected any significant structural degradation caused by the applied thermal–mechanical loads.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 39
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    Report Date: October 1996
    No. Pages: 21
    Funding Organization: WU 505-63-40
    Keywords:      Elevated temperatures; Hypersonic vehicle; Strain gages; Structural analysis; Structural testing; Test techniques
    Notes: Presented at the Structural Testing Technology at High Temperature Conference, Society for Experimental Mechanics, Inc., Dayton, Ohio, Nov. 4–6, 1991. W. Lance Richards, NASA Dryden; Randolph C. Thompson, PRC Inc., Edwards, CA


  33. DESIGN AND INTEGRATION OF AN ACTUATED NOSE STRAKE CONTROL SYSTEM
    Authors: Bradley C. Flick, Michael P. Thomson, Victoria A. Regenie, Keith D. Wichman, Joseph W. Pahle and Michael R. Earls
    Report Number: NASA-TM-104324
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Aircraft flight characteristics at high angles of attack can be improved by controlling vortices shed from the nose. These characteristics have been investigated with the integration of the actuated nose strakes for enhanced rolling (ANSER) control system into the NASA F-18 High Alpha Research Vehicle. Several hardware and software systems were developed to enable performance of the research goals. A strake interface box was developed to perform actuator control and failure detection outside the flight control computer. A three-mode ANSER control law was developed and installed in the Research Flight Control System. The thrust-vectoring mode does not command the strakes. The strakes and thrust-vectoring mode uses a combination of thrust vectoring and strakes for lateral–directional control, and strake mode uses strakes only for lateral–directional control. The system was integrated and tested in the Dryden Flight Research Center (DFRC) simulation for testing before installation in the aircraft. Performance of the ANSER system was monitored in real time during the 89-flight ANSER flight test program in the DFRC Mission Control Center. One discrepancy resulted in a set of research data not being obtained. The experiment was otherwise considered a success with the majority of the research objectives being met.
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    Subject Category: 05
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    Report Date: October 1996
    No. Pages: 30
    Funding Organization: WU 505-68-30
    Keywords:      Ada; Digital systems; F-18 High Alpha Research Vehicle; Flight control; Nose strakes; Servoloop; Servoactuators
    Notes: Presented at the 5th High Angle of Attack Technology Conference, NASA Langley Research Center, Hampton, Virginia, Sept. 17–19, 1996.


  34. EFFECT OF ACTUATED FOREBODY STRAKES ON THE FOREBODY AERODYNAMICS OF THE NASA F-18 HARV , Technical Memorandum
    Authors: David F. Fisher, Daniel G. Murri and Wendy R. Lanser
    Report Number: NASA-TM-4774
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Extensive pressure measurements and off-surface flow visualization were obtained on the forebody and strakes of the NASA F-18 High Alpha Research Vehicle (HARV) equipped with actuated forebody strakes. Forebody yawing moments were obtained by integrating the circumferential pressures on the forebody and strakes. Results show that large yawing moments can be generated with forebody strakes. At angles of attack greater than 40 deg, deflecting one strake at a time resulted in a forebody yawing moment control reversal for small strake deflection angles. At AOA = 40 deg and 50 deg, deflecting the strakes differentially about a 20 deg symmetric strake deployment eliminated the control reversal and produced a near linear variation of forebody yawing moment with differential strake deflection. At AOA = 50 deg and for 0 deg and 20 deg symmetric strake deployments, a larger forebody yawing moment was generated by the forward fuselage (between the radome and the apex of the leading-edge extensions), than on the radome where the actuated forebody strakes were located. Cutouts on the flight vehicle strakes that were not on the wind tunnel models are believed to be responsible for deficits in the suction peaks on the flight radome pressure distributions and differences in the forebody yawing moments.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 02
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    Report Date: October 1996
    No. Pages: 33
    Funding Organization: WU 505-68-30
    Keywords:      Angle-of-attack; F-18 aircraft; Flight control; Flight test; Pressure distribution; Vortices; Yawing moments
    Notes: Presented at the NASA Langley High-Angle-of-Attack Technology Conference, Langley, Sept. 17–19, 1996. D.F. Fisher, NASA DFRC, Edwards, CA, D.G. Murri, NASA Langley, Hampton, VA, W.R. Lanser, Ames, Moffet Field, CA


  35. AN OVERVIEW OF THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE , Technical Memorandum
    Authors: Albion H. Bowers, Joseph W. Pahle, R. Joseph Wilson, Bradley C. Flick and Richard L. Rood
    Report Number: NASA-TM-4772
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper gives an overview of the NASA F-18 High Alpha Research Vehicle. The three flight phases of the program are introduced, along with the specific goals and data examples taken during each phase. The aircraft configuration and systems needed to perform the disciplinary and inter-disciplinary research are discussed. The specific disciplines involved with the flight research are intro-duced, including aerodynamics, controls, propulsion, systems, and structures. Decisions that were made early in the planning of the aircraft project and the results of those decisions are briefly discussed. Each of the three flight phases corresponds to a particular aircraft configuration, and the research dictated the configuration to be flown. The first phase gathered data with the baseline F-18 configuration. The second phase was the thrust-vectoring phase. The third phase used a modified forebody with deployable nose strakes. Aircraft systems supporting these flights included extensive instrumentation systems, integrated research flight controls using flight control hardware and corresponding software, analog interface boxes to control forebody strakes, a thrust-vectoring system using external postexit vanes around axisymmetric nozzles, a forebody vortex control system with strakes, and backup systems using battery-powered emergency systems and a spin recovery parachute.
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    Subject Category: 05
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    Report Date: October 1996
    No. Pages: 39
    Funding Organization: WU 505-68-30
    Keywords:      Aircraft aerodynamics; Control laws; Flight test; High angle-of-attack; Thrust vectoring
    Notes: Presented at the NASA Langley High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia, Sept. 17–19, 1996.


  36. HIGH-ALPHA HANDLING QUALITIES FLIGHT RESEARCH ON THE NASA F/A-18 HIGH ALPHA RESEARCH VEHICLE
    Authors: Keith D. Wichman, Joseph W. Pahle, Catherine Bahm, John B. Davidson, Barton J. Bacon, Patrick C. Murphy, Aaron J. Ostroff and Keith D. Hoffler
    Report Number: NASA-TM-4773
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight research study of high-angle-of-attack handling qualities has been conducted at the NASA Dryden Flight Research Center using the F/A-18 High Alpha Research Vehicle (HARV). The objectives were to create a high-angle-of-attack handling qualities flight database, develop appropriate research evaluation maneuvers, and evaluate high-angle-of-attack handling qualities guidelines and criteria. Using linear and nonlinear simulations and flight research data, the predictions from each criterion were compared with the pilot ratings and comments. Proposed high-angle-of-attack nonlinear design guidelines and proposed handling qualities criteria and guidelines developed using piloted simulation were considered. Recently formulated time-domain Neal-Smith guidelines were also considered for application to high-angle-of-attack maneuvering. Conventional envelope criteria were evaluated for possible extension to the high-angle-of-attack regime. Additionally, the maneuvers were studied as potential evaluation techniques, including a limited validation of the proposed standard evaluation maneuver set. This paper gives an overview of these research objectives through examples and summarizes result highlights. The maneuver development is described briefly, the criteria evaluation is emphasized with example results given, and a brief discussion of the database form and content is presented.
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    Report Date: November 1996
    No. Pages: 37
    Funding Organization: WU 505-68-30
    Keywords:      Criteria guidelines; Database; Handling qualities; High-angle-of-attack; Maneuvers
    Notes: Presented at the NASA Langley High-Angle-of-Attack Conference, Langley Research Center, Sept. 17–19, 1996. Keith Wichman, NASA Dryden Flight Research Center, Edwards, California. Joseph Pahle, NASA Dryden Flight Research Center, Edwards, California


  37. THRUST VECTORING ON THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE
    Authors: Albion H. Bowers and Joseph W. Pahle
    Report Number: NASA-TM-4771
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Investigations into a multiaxis thrust-vectoring system have been conducted on an F-18 configuration. These investigations include ground-based scale-model tests, ground-based full-scale testing, and flight testing. This thrust-vectoring system has been tested on the NASA F-18 High Alpha Research Vehicle (HARV). The system provides thrust vectoring in pitch and yaw axes. Ground-based subscale test data have been gathered as background to the flight phase of the program. Tests investigated aerodynamic interaction and vane control effectiveness. The ground-based full-scale data were gathered from static engine runs with image analysis to determine relative thrust-vectoring effectiveness. Flight tests have been conducted at the NASA Dryden Flight Research Center. Parameter identification input techniques have been developed. Individual vanes were not directly controlled because of a mixer-predictor function built into the flight control laws. Combined effects of the vanes have been measured in flight and compared to combined effects of the vanes as predicted by the coldjet test data. Very good agreement has been found in the linearized effectiveness derivatives.
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    Report Date: November 1996
    No. Pages: 31
    Funding Organization: WU 505-68-30
    Keywords:      Aircraft parameter identification; Cold jet testing; Flight test; Thrust vectoring; Wind tunnel testing
    Notes: Presented at the NASA Langley High-Angle-of-Attack Technology Conference, Langley Research Center, Hampton, Virginia, Sept. 17–19, 1996. Albion H. Bowers and Joseph W. Pahle, NASA Dryden Flight Research Center, Edwards, California.


  38. YEH-STRATTON CRITERION FOR STRESS CONCENTRATIONS ON FIBER-REINFORCED COMPOSITE MATERIALS , Contractor Report
    Authors: Hsien-Yang Yeh and W. Lance Richards
    Report Number: NASA-CR-198054
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This study investigated the Yeh-Stratton Failure Criterion with the stress concentrations on fiber-reinforced composites materials under tensile stresses. The Yeh-Stratton Failure Criterion was developed from the initial yielding of materials based on macromechanics. To investigate this criterion, the influence of the materials anisotropic properties and far field loading on the composite materials with central hole and normal crack were studied. Special emphasis was placed on defining the crack tip stress fields and their applications. The study of Yeh-Stratton criterion for damage zone stress fields on fiber-reinforced composites under tensile loading was compared with several fracture criteria; Tsai-Wu Theory, Hoffman Theory, Fischer Theory, and Cowin Theory. Theoretical predictions from these criteria are examined using experimental results.
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    Subject Category: 24
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    Report Date: November 1996
    No. Pages: 8
    Funding Organization: WU-505-63-84
    Keywords:      Composite materials; Composite research; Failure theory; Macromechanics; Stress concentrations
    Notes: Technical monitor: W. Lance Richards. Hsien-Yang Yeh, Professor, Mechanical Engineering, California State University, Long Beach, California 90840.


  39. 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
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    Report Date: September 1996
    No. Pages: 23
    Funding Organization: WU 505-68-30
    Notes: Previously presented at the NASA High Angle-of-Attack Technology Conference, Inlet Aerodynamics and Propulsion Session, NASA Langley Research Center, September 17–19, 1996.


  40. NEURAL NETWORKS FOR SYSTEM IDENTIFICATION AND ADAPTIVE DYNAMIC INVERSION OF A NONLINEAR MASS-SPRING SYSTEM , Conference Paper
    Authors: Rick Lind and Gary Balas
    Report Number: H-2211
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Multilayer feedforward neural networks can be used to approximate the behavior of nonlinear dynamical systems. System identification is performed using standard back propagation to compute the network weights. Dynamic inversion is an approach to controlling a nonlinear system that inverts the nonlinear dynamics and replaces with a set of desired system dynamics. Neural networks are used to model and invert the nonlinear dynamics. Adaptive dynamic inversion is accomplished by continually training the network to identify the varying system dynamics. A mass-spring system is used to demonstrate the technique.
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    Subject Category: 08
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    Report Date: July 1996
    No. Pages: n.a.
    Funding Organization: 505 63 50
    Keywords:      Adaptive control; Dynamic inversion; Neural network; Nonlinear control; System identification
    Notes: Presented at the International Conference on System Engineering, Las Vegas, Nevada, July 1996. Rick Lind, NASA Dryden Flight Research Center, Edwards, CA; Gary Balas, University of Minnesota, Dept. of Aerospace Engineering, Minneapolis, MN.


  41. OPTIMALLY SCALED H-INFINITY FULL INFORMATION CONTROL SYNTHESIS WITH REAL UNCERTAINTY , AIAA Journal
    Authors: Gary J. Balas, Rick Lind and Andy Packard
    Report Number: H-2212
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An algorithm to synthesize optimal controllers for the scaled H-infinity full information problem with real and complex uncertainty is presented. The control problem is reduced to a linear matrix inequality, which can be solved via a finite dimensional convex optimization. This technique is compared with the optimal scaled H-infinity full information with only complex uncertainty and D-K iteration control design to synthesize controllers for a missile autopilot. Directly including real parametric uncertainty into the control design results in improved robust performance of the missile autopilot. The controller synthesized via D-K iteration achieves results similar to the optimal designs.
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    Subject Category: 08
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    Report Date: July 1996
    No. Pages: n.a.
    Funding Organization: 505 63 50
    Keywords:      Full information; Missile; Parametric uncertainty; Structured singular value
    Notes: Published in the AIAA Journal of Guidance and Dynamics, Vol 19, No. 4, July-August 1996. Gary J. Balas and Rick Lind, University of Minnesota, Minneapolis, Minnesota, Andy Packard, University of California, Berkeley, Berkeley, California.


  42. OPTIMAL FULL INFORMATION SYNTHESIS FOR FLEXIBLE STRUCTURES IMPLEMENTED ON CRAY SUPERCOMPUTERS , AIAA Conference Paper
    Authors: Rick Lind and Gary J. Balas
    Report Number: AIAA-96-3809
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper considers an algorithm for synthesis of optimal controllers for full information feedback. The synthesis procedure reduces to a single linear matrix inequality which may be solved via established convex optimization algorithms. The computational cost of the optimization is investigated. It is demonstrated the problem dimension and corresponding matrices can become large for practical engineering problems. This algorithm represents a process that is impractical for standard workstations for large order systems. A flexible structure is presented as a design example. Control synthesis requires several days on a workstation but may be solved in a reasonable amount of time using a Cray supercomputer.
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    Subject Category: 31
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    Report Date: July 1996
    No. Pages: n.a.
    Funding Organization: 522-32-34
    Keywords:      Flexible structures; Linear matrix inequality; Optimization; Robust control; Supercomputers
    Notes: Presented at AIAA Guidance, Navigation, and Control Conference, San Diego, California, July 1996. Rick Lind, NASA Dryden Flight Research Center; Gary Balas, University of Minnesota.


  43. SUBSPACE IDENTIFICATION WITH MULTIPLE DATA SETS , AIAA Conference
    Authors: Laurent Duchesne, Eric Feron, James D. Paduano and Marty Brenner
    Report Number: AIAA-96-3716
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Most existing subspace identification algorithms assume that a single input to output data set is available. Motivated by a real life problem on the F18-SRA experimental aircraft, we show now these algorithms are readily adapted to handle multiple data sets. We show by means of an example the relevance of such an improvement.
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    Report Date: July 1996
    No. Pages: n.a.
    Funding Organization: 529-50-04
    Keywords:      Aeroelasticity; Modal stability; Structural dynamics; Subspace identification; System identification
    Notes: Presented at AIAA Guidance, Navigation, and Control Conference, San Diego, California, July 1996. Laurent Duchesne, Eric Feron and James Paduano, Massachusetts Institute of Technology; Marty Brenner, NASA Dryden Flight Research Center.


  44. ON SUPERSTABILITY OF SEMIGROUPS , Conference Report
    Authors: A.V. Balakrishnan
    Report Number: H-2302
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents a brief report on superstable semigroups—abstract theory and some applications thereof. The notion of super stability is a strengthening of exponential stability and occurs in Timoshenko models of structures with self-straining material using pure (idealized) rate feedback. It is also relevant to the problem of Riesz bases of eigenfunctions of infinitesimal generators under perturbation.
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    Report Date: July 1996
    No. Pages: 9
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: In Proceedings of Dynamic Systems and Applications, Vol. 5, pp. 371-384, 1996. 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.


  45. ROBUST STABILIZING COMPENSATORS FOR FLEXIBLE STRUCTURES WITH COLLOCATED CONTROLS , Journal Article
    Authors: A.V. Balakrishnan
    Report Number: H-2303
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: For flexible structures with collocated rate and attitude sensors/actuators, we characterize compensator transfer functions which guarantee modal stability even when stiffness/inertia parameters are uncertain. While the compensators are finite-dimensional, the structure models are allowed to be infinite-dimensional (continuum models), with attendance complexity of the notion of stability; thus exponential stability is not possible and the best we can obtain is strong stability. Robustness is interpreted essentially as maintaining stability in the worst case. The conditions require that the compensator transfer functions be positive real and use is made of the Kalman–Yakubovic lemma to characterize them further. The concept of positive realness is shown to be equivalent to dissipativity in infinite dimensions. In particular we show that for a subclass of compensators it is possible to make the system strongly stable as well as dissipative in an appropriate energy norm.
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    Report Date: June 1996
    No. Pages: 26
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      Compensators; Robustness; Positive real; Strong stability; Resolvent; Semigroup; Abstract wave equation.
    Notes: Journal of Applied Mathematics and Optimization, Vol. 33, No. 1, pp. 35-60, 1996. 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.


  46. VIBRATING SYSTEMS WITH SINGULAR MASS-INERTIA MATRICES , Conference Report
    Authors: A.V. Balakrishnan
    Report Number: H-2304
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Vibrating systems with singular mass-inertial matrices arise in recent continuum models of Smart Structures (beams with PZT strips) in assessing the damping attainable with rate feedback. While they do not quite yield “distributed” controls, we shown that they can provide a fixed nonzero lower bound for the damping coefficient at all mode frequencies. The mathematical machinery for modelling the motion involves the theory of Semigroups of Operators. We consider a Timoshenko model for torsion only—a “smart string,” where the damping coefficient turns out to be a constant at all frequencies. We also observe that the damping increases initially with the feedback gain but decreases to zero eventually as the gain increases without limit.
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    Report Date: May 1996
    No. Pages: 8
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: In 1st International Conf. on Nonlinear Problems in Aviation and Aerospace, Daytona Beach, FL, 5-9-96. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  47. IGNITION DELAY ASSOCIATED WITH A STRAINED FUEL STRIP , Conference Report
    Authors: T. J. Gerk and A. R. Karagozian
    Report Number: H-2305
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Ignition processes associated with two adjacent fuel-oxidizer interfaces bounding a strained fuel strip are explored here using single-step activation energy asymptotics. Calculations are made for constant as well as temporally decaying strain fields. Three possible modes of ignition are determined: one in which the two interfaces guide independently as diffusion flames; one in which the two interfaces ignite dependently and in which ignition may be augmented by strain due to enhanced thermal feedback; and one in which ignition occurs to form a single, premixed flame at very high strain rates before ignition is completely prevented. In contrast to a single, isolated interface in which ignition can be prevented by overmatching heat production with heat convection due to strain, ignition of a strained fuel strip can also be prevented if the finite extent of fuel is diluted by oxidizer more quickly than heat production can cause a positive feedback thermal runaway. These behaviors are dependent on the relative sizes of timescales associated with species and heat diffusion, with convection due to strain, and with the chemical reaction. The results here indicate that adjacent, strained species interfaces may ignite quite differently in nature from ignition of a single, strained interface and that their interdependence should be considered as the interfaces are brought closer together in complex strain fields. Critical strain rates leading to complete ignition delay are found to be considerably smaller for the fuel strip than those for single interfaces as the fuel strip is made thin in comparison to diffusion and chemical length scales.
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    Report Date: June 1996
    No. Pages: 8
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: In 26th Symposium (International) on Combustion, pp. 1095-1102. 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.


  48. MODELING AND IN-FLIGHT IMAGING OF TRANSVERSE GAS JETS INJECTED BEHIND A REARWARD-FACING STEP , Conference Report
    Authors: A.R. Karagozian, K.C. Wang, A. -T. Le and O.I. Smith
    Report Number: H-2306
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A model is developed and in-flight experiments are performed to study the behavior of a transverse jet injected behind a rearward-facing step in supersonic flow. The model for jet behavior places emphasis on the dynamics of the vortical structures observed to dominate jet cross-section, as done in prior models. The complex flowfield formed downstream of the step is represented by a combination of empirical and analytical correlations. In-flight experiments to study this flowfield are conducted at NASA Dryden Flight Research Center using a flight test fixture situated under the fuselage of an F-104G aircraft. Model predictions for jet trajectory are compared with prior wind tunnel experiments as well as with results from the present in-flight experiments. Reasonable comparisons are obtained between model and experiments, indicating that this simple phenomenological model could be useful as a design tool in optimizing transverse jet penetration and mixing.
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    Report Date: January 1996
    No. Pages: 7
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: 34th AIAA Aerospace Sciences Meeting, Jan. 1996. Also Journal of Propulsion & Power, V. 12, pp. 1129-1136. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  49. SIMULATION AND VISUALIZATION OF THE FLUID FLOW FIELD IN ANEURYSMS: A VIRTUAL ENVIRONMENTS METHODOLOGY , Conference Report
    Authors: Walter J. Karplus, Michael R. Harreld and Daniel J. Valentino
    Report Number: H-2307
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Virtual environments, of the type described in this paper, employ hardware and software similar to that used for virtual reality displays. However, the viewer is presented with the outputs of computer simulations, involving the solution of systems of nonlinear partial differential equations in three dimensions, superimposed on a real-world physical model. In the present application, the objective is to aid physicians in the performance of a novel and challenging clinical procedure -- the noninvasive treatment of aneurysms located deep within the brain. To this end, the physician is given the opportunity to “taking a walk” inside the aneurysm and to observe the transient pressure and stress fields in the blood flowing in the aneurysm, as the heart pulses and as therapeutic procedures are implemented.
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    Report Date: October 1996
    No. Pages: 7
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Proc. of 7th Intl. Symp. on Micro Machine & Human Science, Japan, IEEE & Nagoya Industrial Res. Ins., p. 25-31. Work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  50. THE VAN CAPELLE AND DURRER MODEL OF CARDIAC ACTION POTENTIAL GENERATION AND 2D PROPAGATION: MODIFICATIONS AND APPLICATION TO SPIRAL WAVE PROPAGATION , Conference Report
    Authors: Boris Y. Kogan, Walter J. Karplus and Mikhail G. Karpoukhin
    Report Number: H-2308
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The properties of the Van Capelle and Durrer (VC-D) simplified model are studies to estimate how adequately it reproduces the physiological properties of cardiac cells and tissue. It is shown that the transition from a membrane capacitance of 10 mu F/cm(2), used in previous studies, to 1 mu F/cm(2), the introduction of adjustable action potential duration restitution, and the appropriate adjustment of action potential duration (APD) bring the simulation result closer to reality. In particular, it is shown that in the 2D model of anisotropic tissue, nonstationary propagation of spiral waves in response to premature beat is possible when the cells of the tissue have specified APD restitution properties. The continuous pacemaker activity in this case leads to the creation of multiple spiral waves.
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    Report Date: January 1996
    No. Pages: 7
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      Van Capelle and Durrer simplified model; 2D spiral waves; action potential duration restitution; simulation on parallel computers
    Notes: Proc. of 1996 Western Multiconference: Simulation in the Medical Sciences. Soc. for Computer Simulation, pp. 106-112. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Cont. monitor, R. Quinn; funding facil.by K. Iliff


  51. PROPAGATION OF ELECTRICAL EXCITATION IN A RING OF CARDIAC CELLS: A COMPUTER SIMULATION STUDY , Conference Report
    Authors: B. Y. Kogan, W. J. Karplus, M. G. Karpoukhin, I.M. Roizen, E. Chudin and Z. Qu
    Report Number: H-2309
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The propagation of electrical excitation in a ring of cells described by the Noble, Beeler-Reuter (BR), Luo-Rudy I (LR I), and third-order simplified (TOS) mathematical models is studied using computer simulation. For each of the models it is shown that after transition from steady-state circulation to quasi-periodicity achieved by shortening the ring length (RL), the action potential duration (APD) restitution curve becomes a double-valued function and is located below the original (that of an isolated cell) APD restitution curve. The distributed of APD and diastolic interval (DI) along a ring for the entire range of RL corresponding to quasi-periodic oscillations remain periodic with the period slightly different from two RLs. The “S” shape of the original APD restitution curve determines the appearance of the second steady-state circulation region for short RLs. For all the models and the wide variety of their original APD restitution curves, no transition from quasi-periodicity to chaos was observed
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    Report Date: June 1996
    No. Pages: 11
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: In Simulation Modelling in Bioengineering, pp. 303-313, Boston Computational Mechanics Publications, 1996. Work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  52. MIXING ENHANCEMENT IN A LOBED INJECTOR , UCLA Masters thesis
    Authors: L.L. Smith, A.J. Majamaki, I.T. Lam, O. Delabroy, A.R. Karagozian, F.E. Marble and O.I. Smith
    Report Number: H-2310
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An experimental investigation of the non-reactive mixing processes associated with a lobed fuel injector in a coflowing air stream is presented. The lobed fuel injector is a device which generates streamwise vorticity, producing high strain rates which can enhance the mixing of reactants while delaying ignition in a controlled manner. The lobed injectors examined in the present study consist of two corrugated plates between which a fuel surrogate CO(2), is injected into coflowing air. Acetone is seeded in the CO(2) supply as a fuel marker. Comparison of two alternative lobed injector geometries is made with a straight fuel injector to determine net differences in mixing and strain fields due to streamwise vorticity generation. Planar laser-induced fluorescence (PLIF) of the seeded acetone yields two-dimensional images of the scalar concentration field at various downstream locations, from which local mixing and scalar dissipation rates are computed. It is found that the lobed injector geometry can enhance molecular mixing and create a highly strained flowfield, and that the strain rates generated by scalar energy dissipation can potentially delay ignition in a reacting flowfield.
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    Report Date: July 1996
    No. Pages: 12
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: M.Sc. thesis, UCLA, 1996. 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.


  53. ENHANCEMENT OF NUCLEATE BOILING HEAT FLUX ON MACRO/MICRO-STRUCTURED SURFACES COOLED BY MULTIPLE IMPINGING JETS , Conference Report
    Authors: S. Kugler and V.K. Dhir
    Report Number: H-2311
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An experimental investigation of nucleate boiling heat transfer from modified surfaces cooled by multiple in-line impinging circular jets is reported and found to agree with single jet results. To increase the maximum heat flux (CHF) and heat removal rate, the boiling surface was modified by both macro and micro enhancements. Macro modification consisted of machined radial grooves in the boiling surface, while micro modification was achieved by sintering. Macro enhancements allow higher CHF at lower superheat, with both enhancements pushing the nucleate boiling curve upwards and to the left. With the addition of both micro and macro structured enhancements, maximum heat flux and nucleate boiling can be enhanced by more than 200 percent.
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    Report Date: June 1996
    No. Pages: n.a.
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Heat Transfer, Houston, 1996. AIChE Symposium Series, Vol. 92, pp. 287-293, American Inst. of Chemical Engineers. Work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  54. NOX EMISSIONS FROM A LOBED FUEL INJECTOR/BURNER , Conference Report
    Authors: M.G. Mitchell, L.L. Smith, A.R. Karagozian and O.I. Smith
    Report Number: H-2312
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The present experimental study examines the performance of a novel fuel injector/burner configuration with respect to reduction in nitrogen oxide NOx emissions. The lobed injector/burner is a device in which very rapid initial mixing of reactants can occur through strong streamwise vorticity generation, producing high fluid mechanical strain rates which can delay ignition and thus prevent the formation of stoichiometric diffusion flames. Further downstream of the rapid mixing region, this flowfield produces a reduced effective strain rate, thus allowing ignition to occur in a premixed mode, where it is possible for combustion to take place under locally lean conditions, potentially reducing NOx emissions from the burner. The present experiments compare NO/NO2/NOx emissions from a lobed fuel injector configuration with emissions from a straight fuel injector to determine the net effect of streamwise vorticity generation. Preliminary results show that the lobed injector geometry can produce lean premixed flame structures, while for comparable flow conditions, a straight fuel injector geometry produces much longer, sooting diffusion flames or slightly rich premixed flames. NOx measurements show that emissions from a lobed fuel injector/burner can be made significantly lower than from a straight fuel injector under comparable flow conditions.
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    Report Date: October 1996
    No. Pages: 16
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Presented at Western States Section/The Combustion Institute Fall Meeting, 1996. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  55. EFFECT OF DENSITY GRADIENTS IN CONFINED SUPERSONIC SHEAR LAYERS. I. TWO-DIMENSIONAL DISTURBANCES , Conference Report
    Authors: Oshin Peroomian and R. E. Kelly
    Report Number: H-2313
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The effect of density gradients on two-dimensional (2-D) supersonic wall modes (acoustic modes) of a 2-D confined compressible shear layer are investigated using linear analysis. Due to the inadequacies of the hyperbolic tangent velocity profile, the boundary layer basic flow profiles are used. First, a test case is taken with the same parameters as in a previous analysis by Tam and Hu [J. Fluid Mech. 203, 51 (1989)], who used a hyperbolic tangent profile. For the boundary layer profiles, three generalized inflection points are found giving rise to three modes. The first two show similar properties to those found by Tam and Hu, whereas the third is a new mode that can have a higher growth rate than the others. As the density ratio is increased above that of the test case, the smallest of the three neutral phase speeds tends toward the speed of the lower-velocity stream, and the other two eventually coalesce and then disappear. These two effects lead to a linear resonance between some of the modes that increases the cutoff frequency and growth rate of the lowest mode. In face, growth rates of two to four times the test case were found as the density ratio was increased to 7. A similar trend (linear resonance) is observed as the density ratio is decreased from the test case, but the growth rate is only slightly changed from the test case.
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    Report Date: January 1996
    No. Pages: 16
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Phys. Fluids, Vol 8, No. 1, pp. 225-240, January, 1996. 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.


  56. EFFECT OF DENSITY GRADIENTS IN CONFINED SUPERSONIC SHEAR LAYERS. II. THREE-DIMENSIONAL MODES , Conference Report
    Authors: Oshin Peroomian
    Report Number: H-2314
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The effect of density gradients on three-dimensional (3-D) supersonic acoustic modes and 3-D subsonic modes in a compressible confined shear layer were investigated using linear analysis. The compressible Rayleigh equation with the solution of the steady compressible boundary layer equations as its basic profiles was solved for two different density ratios, 1.398 and 3.0, and several spanwise wave numbers. For a density ratio of 1.398, the 2-D C(01) mode had the highest growth rate for a fixed aspect ratio B/H = 2. At the higher-density ratio, and 3-D mode had a slightly higher maximum growth rate than the two-dimensional (2-D) mode, and the maximum growth rate was at a location where the phase speed was supersonic with respect to the fact stream.
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    Report Date: January 1996
    No. Pages: 7
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Phys. Fluids, Vol. 8, No. 1, pp. 241-247, January 1996. 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.


  57. A FEASIBILITY STUDY OF A CELL AVERAGE BASED MULTI-DIMENSIONAL ENO SCHEME FOR USE IN SUPERSONIC SHEAR LAYERS , Conference Report
    Authors: Oshin Peroomian and Sukumar Chakravarthy
    Report Number: H-2315
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The performance of a multi-dimensional cell average based ENO scheme is evaluated for use in supersonic confined shear layers. Different ENO recipes for limiting the polynomial coefficients are investigated and their results documented. For shear layers with strong shocks a coefficient-by-coefficient technique must be utilized in order to avoid strong oscillations near discontinuities. Also, ENO polynomial interpolation on primitive variables, calculated from the conservation variables was studies. Oscillations in pressure and species concentration were greatly reduced by this process.
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    Report Date: January 1996
    No. Pages: 12
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: AIAA Paper No. 96-0524. At 34th Aerospace Sciences Meeting & Exhibit, Reno, NV, January 15-18, 1996. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  58. SPATIAL SIMULATIONS OF A CONFINED SUPERSONIC SHEAR LAYER AT TWO DENSITY RATIOS , Conference Report
    Authors: Oshin Peroomian, R. E. Kelly and Sukumar Chakravarthy
    Report Number: H-2316
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Spatial 2-D simulations of a confined shear layer are carried out for two density ratios using a second order TVD scheme. The density ratios are chosen based on the linear stability results of Peroomian and Kelly[1]. For both sets of parameters, only the acoustic instability modes exist. Solutions of the steady compressible boundary layer equations which are taken to be functions of the cross-stream direction only are used to initialize the computational box. For the first case, formation of strong shock-expansion structures are observed within the flowfield. The acceleration and rotation of the lower part of these structures due to the generation of vorticity near the walls increase the magnitude of the pressure change near the lower wall, causing a blowing effect from the wall which improves the entrainment process. For the higher density ratio, some interesting phenomena such as generation of Kelvin-Helmholtz type structures in the subsonic portion and compression-expansion waves in the supersonic portion of the shear layer were observed. Also, the fundamental mode was observed to saturate quickly and then, after a long plateau, was observed to grow again.
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    Report Date: January 1996
    No. Pages: 19
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: AIAA Paper No. 96-0783. At 34th Aerospace Sciences Meeting & Exhibit, Reno, NV, January 15-18, 1996. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  59. IGNITION, BURNING, AND EXTINCTION OF A STRAINED FUEL STRIP , Conference Report
    Authors: T. Selerland and A. R. Karagozian
    Report Number: H-2317
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flame structure and ignition and extinction processes associated with a strained fuel strip are explored numerically using detailed transport and complex kinetics for a propane-air reaction. Ignition modes are identified that are similar to those predicted by one-step activation energy asympototics, i.e., modes in which diffusion flames can ignite as independent or dependent interfaces, and modes in which single premixed or partially premixed flames ignite. These ignition modes are found to be dependent on critical combinations of strain rate, fuel strip thickness, and initial reactant temperatures. Extinction of this configuration is seen to occur due to fuel consumption by adjacent flames, although viscosity is seen to have the effect of delaying extinction by reducing the effective strain rate and velocity field experienced by the flames.
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    Report Date: October 1996
    No. Pages: 17
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Presented at Western States Section/The Combustion Institute Fall Meeting, 1996. 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.


  60. ATMOSPHERIC LEE WAVES , Journal Article
    Authors: M. G. Wurtele, R. D. Sharman and A. Datta
    Report Number: H-2318
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The atmospheric lee wave is a disturbance propagated by buoyancy and arising from an isolated source, usually by flow over ridges and mountains. Part of this review treats two-dimensional solutions, both Boussinesq and non-Boussinesq, linear and nonlinear. These discussions emphasize trapped waves, the downslope windstorm, the drag on the earth and the upward momentum flux, the hydro-static approximation and its limitations, effects of critical layers, and middle atmospheric wave breaking. Three-dimensional Boussinesq linear and nonlinear solutions are also discussed; shown are the variety of regimes possible, from ship waves to shedding vortices. Photographs of natural phenomena are presented as realizations, together with relevant numerical simulation graphics. The difficulties and achievements of simulation models are also outlined.
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    Report Date: December 1996
    No. Pages: 48
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      Mountain waves; Gravity waves; Mesoscale phenomena
    Notes: Annual Rev. Fluid. Mech., Vol. 28, pp. 429-476, 1996. 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.


  61. THE PROPAGATION OF GRAVITY–INERTIA WAVES AND LEE WAVES UNDER A CRITICAL LEVEL , Journal Article
    Authors: M. G. Wurtele, A. Datta and R. D. Sharman
    Report Number: H-2319
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: As is well known, the linear dynamic equations for gravity–inertia waves are characterized by three singular levels, one being the critical level at which flow speed and wave speed are equal, and the other two at which the flow speed is equal to ± flk, where f is the Coriolis parameter and k the wave number, herein called Rossby singularities. This article discusses the propagation of two-dimensional gravity–inertial disturbances, both monochromatic and with continuous spectrum (i.e., lee waves), in a direction toward all of these singular levels. The study is conducted by analysis, which provides closed-form solutions to the linear equations, and by numerical simulation, which confirms the analysis and also exhibits nonlinearities where these are significant. It is found that the Rossby singularity produces nonlinear reflection of a monochromatic wave, and comparisons are made with the case of the pure gravity wave (f = 0) reflected by a critical level. Unlike that situation, in the present problem the momentum flux is also singular at the reflecting level. However, this is no longer the case when the disturbance contains a continuous spectrum, as in a lee wave produced by a smooth isolated ridge. In this case, the problem is essentially linear, and a relatively simple analytic approximation to the solution is presented and verified by simulation. The critical level acts as a lid but produces no singular effects. However, certain types of forcing profiles are identified that, despite being themselves of small amplitude, do in fact lead to nonlinearities in the field.
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    Report Date: June 1996
    No. Pages: 19
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Journal of the Atmospheric Sciences, Vol. 53, No. 11, 1 June 1996. 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.


  62. DYNAMICS AND CONTROL OF ARTICULATED ANISOTROPIC TIMOSHENKO BEAMS , Technical Report
    Authors: A. V. Balakrishnan
    Report Number: H-2320
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper illustrates the use of continuum models in control design for stabilizing flexible structures. A 6-DOF anisotropic Timoshenko beam with discrete nodes where lumped masses or actuators are located provides a sufficiently rich model to be of interest for mathematical theory as well as practical application. We developed concepts and tools to help answer engineering questions without having to resort to ad hoc heuristic (“physical”) arguments or faith. In this sense the paper is more mathematically oriented than engineering papers and vice versa at the same time. For instance we make precise time-domain solutions using the theory of semigroups of operators rather than formal “inverse Laplace transforms.” We show that the modes arise as eigenvalues of the generator of the semigroup, which are then related to the eigenvalues of the stiffness operator. With the feedback control, the modes are no longer orthogonal and the question naturally arises as to whether there is still a modal expansion. Here we prove that the eigenfunctions yield a biorthogonal Riesz basis and indicate the corresponding expansion. We prove mathematically that the number of eigenvalues is nonfinite, based on the theory of zeros of entire functions. We make precise the notion of asymptotic modes and indicate how to calculate them. Although limited by space, we do consider the root locus problem and show for instance that the damping at first increases as the control gain increases but starts to decrease at a critical value, and goes to zero as the gain increases without bound. The undamped oscillatory modes remain oscillatory and the rigid-body modes go over into deadbeat modes. The Timoshenko model dynamics are translated into a cononical wave equation in a Hilbert space. The solution is shown to require the use of an “energy” norm which is no more than the total energy: potential plus kinetic. We show that, under an appropriate extension of the notion of controllability, rate feedback with a collocated sensor can stabilize the structure in the sense that all modes are damped and the energy decays to zero. An example, non-numeric, is worked out in some detail illustrating the concepts and theory developed.
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    Report Date: September 1996
    No. Pages: 91
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Technical Report No. 1-FSRC-96, September 1996. Dynamics and Control of Distributed Systems, to appear 1998. Result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff