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  1. SELECTED EXAMPLES OF NACA/NASA SUPERSONIC FLIGHT RESEARCH , Special Publication
    Authors: Edwin J. Saltzman and Theodore G. Ayers
    Report Number: NASA-SP-513
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The present Dryden Flight Research Center, a part of the National Aeronautics and Space Administration, has a flight research history that extends back to the mid-1940's. The parent organization was a part of the National Advisory Committee for Aeronautics and was formed in 1946 as the Muroc Flight Test Unit. This document describes 13 selected examples of important supersonic flight research conducted from the Mojave Desert location of the Dryden Flight Research Center over a 4 decade period beginning in 1946. The research described herein was either obtained at supersonic speeds or enabled subsequent aircraft to penetrate or traverse the supersonic region. In some instances there accrued from these research efforts benefits which are also applicable at lower or higher speed regions. A major consideration in the selection of the various research topics was the lasting impact they have had, or will have, on subsequent supersonic flight vehicle design, efficiency, safety, and performance or upon improved supersonic research techniques.
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    Subject Category: 99
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    Report Date: January 1994
    Keywords:      Area rule; Compressibility effects; Digital fly-by-wire; Flight path control; Inertial coupling; Model-to-full scale correlation; Parameter estimation; Reaction controls; Supersonic flight research/testing


  2. ON THE ESTIMATION ALGORITHM FOR ADAPTIVE PERFORMANCE OPTIMIZATION OF TURBOFAN ENGINES
    Authors: Martin Espana and Glenn B. Gilyard
    Report Number: NASA-TM-4551
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The performance seeking control algorithm is designed to continuously optimize the performance of propulsion systems. The performance seeking control algorithm uses a nominal model of the propulsion system and estimates, in flight, the engine deviation parameters characterizing the engine deviations with respect to nominal conditions. In practice, because of measurement biases and/or model uncertainties, the estimated engine deviation parameters may not reflect the engine's actual off-nominal condition. This factor has a necessary impact on the overall performance seeking control scheme exacerbated by the open-loop character of the algorithm. The effects produced by unknown measurement biases over the estimation algorithm are evaluated. This evaluation allows for identification of the most critical measurements for application of the performance seeking control algorithm to an F100 engine. An equivalence relation between the biases and engine deviation parameters stems from an observability study; therefore, it is undecided whether the estimated engine deviation parameters represent the actual engine deviation or whether they simply reflect the measurement biases. A new algorithm, based on the engine's (steady-state) optimization model, is proposed and tested with flight data. When compared with previous Kalman filter schemes, based on local engine dynamic models, the new algorithm is easier to design and tune and it reduces the computational burden of the onboard computer.
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    Report Date: January 1994
    Keywords:      Turbofan Engines; Adaptive Control; Control Systems Design; Optimal Control; Propulsion System Configurations; Propulsion System Performance
    Notes: Also presented as AIAA Paper 93-1823


  3. IN-FLIGHT LIFT-DRAG CHARACTERISTICS FOR A FORWARD-SWEPT-WING AIRCRAFT (AND COMPARISIONS WITH CONTEMPORARY AIRCRAFT) , Technical Paper
    Authors: Edwin J. Saltzman and John W. Hicks
    Report Number: NASA-TP-3414
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Lift (L) and drag (D) characteristics have been obtained in flight for the X-29A airplane (a forward-swept-wing demonstrator) for Mach numbers (M) from 0.4 to 1.3. Most of the data were obtained near an altitude of 30,000 ft. A representative Reynolds number, for M = 0.9 and a pressure altitude of 30,000 ft, is based on the mean aerodynamic chord. The X-29A data (forward-swept wing) are compared with three high-performance fighter aircraft—the F-15C, F-16C, and F/A18. The lifting efficiency of the X-29A, as defined by the Oswald lifting efficiency factor, e, is about average for a cantilevered monoplane for M = 0.6 and angles of attack up to those required for maximum L/D. At M = 0.6 the level of L/D and e, as a function of load factor, for the X-29A was about the same as for the contemporary aircraft. The X-29A and its contemporaries have high tran-sonic wave drag and equivalent parasite area compared with aircraft of the 1940s through 1960s.
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    Report Date: December 1994
    No. Pages: 58
    Funding Organization: WU 505-68-50
    Keywords:      Efficiency; Forward-swept wing; Lift-drag; Lift-related drag; Aerodynamic cleanners; Transonic wave drag


  4. FLOW-VISUALIZATION STUDY OF THE X-29A AIRCRAFT AT HIGH ANGLE OF ATTACK USING A 1/48-SCALE MODEL
    Authors: Stacey J. Cotton and Lisa J. Bjarke
    Report Number: NASA-TM-104268
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A water-tunnel study on a 1/48-scale model of the X-29A aircraft was performed at the NASA Dryden Flow Visualization Facility. The water-tunnel test enhanced the results of the X-29A flight tests by providing flow-visualization data for comparison and insights into the aerodynamic characteristics of the aircraft. The model was placed in the water tunnel at angles of attack from 20 deg to 55 deg and with angles of sideslip from 0 deg to 5 deg. In general, flow-visualization techniques provided useful information on vortex formation, separation, and breakdown and their role in yaw asymmetries and tail buffeting. Asymmetric forebody vortices were observed at angles of attack greater than 30 deg with 0 deg sideslip and greater than 20 deg with 5 deg sideslip. While the asymmetric flows observed in the water tunnel did not agree fully with the flight data, they did show some of the same trends. In addition, the flow visualization indicated that the interaction of forebody vortices and the wing wake at angles of attack between 20 deg and 35 deg may cause vertical-tail buffeting observed in flight.
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    Report Date: August 1994
    No. Pages: 49
    Funding Organization: WU 533-02-38
    Keywords:      Flow visualization; High angle of attack; Water tunnel; X-29A


  5. MECHANICAL AND THERMAL BUCKLING ANALYSIS OF RECTANGULAR SANDWICH PANELS UNDER DIFFERENT EDGE CONDITIONS , NASA Technical Memorandum
    Authors: William L. Ko
    Report Number: NASA-TM-4585
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The combined load (mechanical or thermal load) buckling equations were established for orthotropic rectangular sandwich panels under four different edge conditions by using the Rayleigh-Ritz method of minimizing the total potential energy of a structural system. Two-dimensional buckling interaction curves and three-dimensional buckling interaction surfaces were constructed for high- temperature honeycomb-core sandwich panels supported under four different edge conditions. The interaction surfaces provide overall comparison of the panel buckling strengths and the domains of symmetrical and antisymmetrical buckling associated with the different edge conditions. In addition, thermal buckling curves of these sandwich panels are presented. The thermal buckling conditions for the cases with and without thermal moments were found to be identical for the small deformation theory.
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    Subject Category: 39
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    Report Date: April 1994
    No. Pages: 67
    Funding Organization: RTOP 532-09-01
    Keywords:      Different edge conditions; Mechanical buckling; Rayleigh-Ritz method; Sandwich panels; Thermal buckling
    Notes: n.a.


  6. DEVELOPING AND FLIGHT TESTING THE HL-10 LIFTING BODY: A PRECURSOR TO THE SPACE SHUTTLE , Reference Publication
    Authors: Robert W. Kempel, Weneth D. Painter and Milton O. Thompson
    Report Number: NASA-RP-1332
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The origins of the lifting-body idea are traced back to the mid-1950s, when the concept of a manned satellite reentering the Earth's atmosphere in the form of a wingless lifting body was first proposed. The advantages of low reentry deceleration loads, range capability, and horizontal landing of a lifting reentry vehicle (as compared with the high deceleration loads and parachute landing of a capsule) are presented. The evolution of the hypersonic HL-10 lifting body is reviewed from the theoretical design and development process to its selection as one of two low-speed flight vehicles for fabrication and piloted flight testing. The design, development, and flight testing of the low-speed, air-launched, rocket-powered HL-10 was part of an unprecedented NASA and contractor effort. NASA Langley Research Center conceived and developed the vehicle shape and conducted numerous theoretical, experimental, and wind-tunnel studies. NASA Flight Research Center (now NASA Dryden Flight Research Center) was responsible for final low-speed (Mach numbers less than 2.0) aerodynamic analysis, piloted simulation, control law development, and flight tests. The prime contractor, Northrop Corp., was responsible for hardware design, fabrication, and integration. Interesting and unusual events in the flight testing are presented with a review of significant problems encountered in the first flight and how they were solved. Impressions by the pilots who flew the HL-10 are included. The HL-10 completed a successful 37-flight program, achieved the highest Mach number and altitude of this class vehicle, and contributed to the technology base used to develop the space shuttle and future generations of lifting bodies.
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    Report Date: April 1994
    No. Pages: 57
    Funding Organization: WU 505-68-50
    Keywords:      HL-10; Lifting body; Rocket airplanes; Wingless flight
    Notes: Mr. Kempel is currently employed by PRC Inc., Edwards, CA. Mr. Painter is affiliated with the National Test Pilot School, Mojave, CA. Mr. Thompson (1926-1993) was employed at NASA Dryden upon his death.


  7. THERMOCRYOGENIC BUCKLING AND STRESS ANALYSES OF A PARTIALLY FILLED CRYOGENIC TANK SUBJECTED TO CYLINDRICAL STRIP HEATING , Technical Memorandum
    Authors: William L. Ko
    Report Number: NASA-TM-4579
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Thermocryogenic buckling and stress analyses were conducted on a horizontally oriented cryogenic tank using the finite element method. The tank is a finite-length circular cylindrical shell with its two ends capped with hemispherical shells. The tank is subjected to cylindrical strip heating in the region above the liquid-cryogen fill level and to cryogenic cooling below the fill level (i.e., under thermocryogenic loading). The effects of cryogen fill level on the buckling temperature and thermocryogenic stress field were investigated in detail. Both the buckling temperature and stress magnitudes were relatively insensitive to the cryogen fill level. The buckling temperature, however, was quite sensitive to the radius-to-thickness ratio. A mechanical stress analysis of the tank also was conducted when the tank was under (1) cryogen liquid pressure loading, (2) internal pressure loading and (3) tank-wall inertia loading. Deformed shapes of the cryogenic tanks under different loading conditions were shown, and high-stress domains were mapped on the tank wall for the strain-gage installations. The accuracies of solutions from different finite element models were compared.
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    Subject Category: 39
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    Report Date: November 1994
    No. Pages: 68
    Funding Organization: WU 505-70-63
    Keywords:      Cryogenic tank; Thermocryogenic buckling; Stress analysis; Thermocryogenic loading; Internal pressure loading; Liquid pressure loading; Tank wall inertia loading


  8. STRAIN-GAGE SELECTION IN LOADS EQUATIONS USING A GENETIC ALGORITHM , NASA Contractor Report
    Authors: Sigurd A. Nelson II
    Report Number: NASA-CR-4597
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Traditionally, structural loads are measured using strain gages. A loads calibration test must be done before loads can be accurately measured. In one measurement method, a series of point loads is applied to the structure, and loads equations are derived via the least squares curve fitting algorithm using the strain gage responses to the applied point loads. However, many research structures are highly instrumented with strain gages, and the number and selection of gages used in a loads equation can be problematic. This paper presents an improved technique using a genetic algorithm to choose the strain gages used in the loads equations. Also presented are a comparison of the genetic algorithm performance with the current T-value technique and a variant known as the Best Step-down technique. Examples are shown using aerospace vehicle wings of high and low aspect ratio. In addition, a significant limitation in the current methods is revealed. The genetic algorithm arrived at a comparable or superior set of gages with significantly less human effort, and could be applied in instances when the current methods could not.
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    Report Date: October 1994
    No. Pages: 20
    Funding Organization: NAS2-13445
    Keywords:      Columns space; Data channel selection; Gage selection methods; Genetic algorithms; Loads equations; Structural calibration
    Notes: NASA Dryden Technical Monitor: Steve Thornton


  9. AIRCRAFT GROUND VIBRATION TESTING AT THE NASA DRYDEN RESEARCH FACILITY - 1993 , NASA Technical Memorandum
    Authors: Michael W. Kehoe and Lawrence C. Freudinger
    Report Number: NASA-TM-104275
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA Dryden Flight Research Facility performs ground vibration testing to assess the structural characteristics of new and modified research vehicles. This paper updates the research activities, techniques used, and experiences in applying this technology to aircraft since 1987. Test equipment, data analysis methods, and test procedures used for typical test programs are discussed. The data presented illustrate the use of modal test and analysis in flight research programs for a variety of aircraft. This includes a technique to acquire control surface free-play measurements on the X-31 airplane more efficiently, and to assess the effects of structural modifications on the modal characteristics of an F-18 aircraft. In addition, the status and results from current research activities are presented. These data show the effectiveness of the discrete modal filter as a preprocessor to uncouple response measurements into simple single-degree-of-freedom responses, a database for the comparison of different excitation methods on a JetStar airplane, and the effect of heating on modal frequency and damping.
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    Report Date: April 1994
    No. Pages: 23
    Funding Organization: RTOP 505-63-50
    Keywords:      Ground vibration test; Hot structures; Modal analysis; Modal filter
    Notes: Also presented at the 12th International Modal Analysis Conference (IMAC), Honolulu, HI, Feb. 2, 1994. Sponsored by SEM.


  10. PRESSURE-SENSING PERFORMANCE OF UPRIGHT CYLINDERS IN A MACH 10 BOUNDARY-LAYER , Technical Memorandum
    Authors: Steven Johnson and Kelly Murphy
    Report Number: NASA-TM-4633
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An experimental research program to provide basic knowledge of the pressure-sensing performance of upright, flush-ported cylinders in a hypersonic boundary layer is described. Three upright cylinders of 0.25-, 0.5-, and 1-in. diameters and a conventional rake were placed in the test section sidewall boundary layer of the 31 Inch Mach 10 Wind Tunnel at NASA Langley Research Center, Hampton, Virginia. Boundary-layer pressures from these cylinders were compared to those measured with a conventional rake. A boundary-layer thickness-to-cylinder-diameter ratio of 8 proved sufficient to accurately measure an overall pressure profile and ascertain the boundarylayer thickness. Effects of Reynolds number, flow angularity, and shock wave impingement on pressure measurement were also investigated. Although Reynolds number effects were negligible at the conditions studied, flow angularity above 10 deg significantly affects the measured pressures. Shock wave impingement was used to investigate orifice-to-orifice pressure crosstalk. No crosstalk was measured. The lower pressure measured above the oblique shock wave impingement showed no influence of the higher pressure generated at the lower port locations.
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    Subject Category: 02
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    Report Date: July 1994
    No. Pages: 25
    Funding Organization: WU 505-70
    Keywords:      Boundary-layer measurements; Flow angularity; Hypersonic aerodynamics; Shock wave impingements; Upright cylinders; Wind tunnel tests
    Notes: Kelly Murphy is affiliated with NASA Langley Research Center, Hampton, Virginia 23665-5225.


  11. SUPERSONIC FLIGHT TEST RESULTS OF A PERFORMANCE SEEKING CONTROL ALGORITHM ON A NASA F-15 AIRCRAFT , AIAA Conference Paper
    Authors: John S. Orme and Timothy R. Conners
    Report Number: AIAA-94-3210
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A model-based, adaptive control algorithm called Performance Seeking Control (PSC) has been flight tested on an F-15 aircraft. The algorithm attempts to optimize performance of the integrated propulsion system during steady-state engine operation. The final phase of a 3-year PSC flight test program is described in this paper. Previous studies of use of PSC on the F-15 airplane show improvement in propulsion system performance. Because these studies were conducted using one of two F-15 engines, the full effect on aircraft performance was not measured. During the most recent studies, both engines were optimized to demonstrate the full effect of PSC propulsion system optimization on aircraft performance. Results were gathered over the 1-g supersonic envelope demonstrating benefits of the integrated control approach. Quantitative flight results illustrating the PSC method for deriving benefits from the F-15 integrated propulsion system for Mach numbers up to 2 are also presented.
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    Report Date: June 1994
    No. Pages: 20
    Funding Organization: n.a.
    Keywords:      n.a.
    Notes: 30th AIAA/ASME/SAE/ASEE Joint Propulsion Conference June 27-29, 1994 / Indianapolis, IN


  12. FURTHER DEVELOPMENT AND FLIGHT TEST OF AN AUTONOMOUS PRECISION LANDING SYSTEM USING A PARAFOIL , Technical Memorandum
    Authors: James E. Murray, Alex G. Sim, David C. Neufeld, Patrick K. Rennich, Stephen R. Norris and Wesley S. Hughes
    Report Number: NASA-TM-4599
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA Dryden Flight Research Center and NASA Johnson Space Center are jointly conducting a phased program to determine the feasibility of the autonomous recovery of a spacecraft using a ram-air parafoil system for the final stages of entry from space to a precision landing. The feasibility is being studied using a flight model of a spacecraft in the generic shape of a flattened biconic that weighs approximately 120 lb and is flown under a commercially available ram-air parafoil. Key components of the vehicle include the Global Positioning System (GPS) guidance for navigation, a flight control computer, an electronic compass, a yaw rate gyro, and an onboard data recorder. A flight test program is being used to develop and refine the vehicle. The primary flight goal is to demonstrate autonomous flight from an altitude of 3,000 m (10,000 ft) with a lateral offset of 1.6 km (1.0 mi) to a precision soft landing. This paper summarizes the progress to date. Much of the navigation system has been tested, including a heading tracker that was developed using parameter estimation techniques and a complementary filter. The autoland portion of the autopilot is still in development. The feasibility of conducting the flare maneuver without servoactuators was investigated as a means of significantly reducing the servoactuator rate and load requirements.
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    Report Date: July 1994
    No. Pages: 13
    Funding Organization: WU 505-68-50
    Keywords:      Aeroelasticity; Flight flutter testing; Ground vibration testing
    Notes: This was prepared as AIAA-94-2141 for the 6th Biennial Flight Test Conference, Colorado Springs, CO, June 20-23, 1994.


  13. FLIGHT TESTING A PROPULSION-CONTROLLED AIRCRAFT EMERGENCY FLIGHT CONTROL SYSTEM ON AN F-15 AIRPLANE
    Authors: F. W. Burcham, Jr., John Burken and Trindel A. Maine
    Report Number: NASA-TM-4590
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests of a propulsion-controlled aircraft (PCA) system on an F-15 airplane have been conducted at the NASA Dryden Flight Research Center. The airplane was flown with all flight control surfaces locked both in the manual throttles-only mode and in an augmented system mode. In the latter mode, pilot thumbwheel commands and aircraft feedback parameters were used to position the throttles. Flight evaluation results showed that the PCA system can be used to land an airplane that has suffered a major flight control system failure safely. The PCA system was used to recover the F-15 airplane from a severe upset condition, descend, and land. Pilots from NASA, U.S. Air Force, U.S. Navy, and McDonnell Douglas Aerospace evaluated the PCA system and were favorably impressed with its capability. Manual throttles-only approaches were unsuccessful. This paper describes the PCA system operation and testing. It also presents flight test results and pilot comments.
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    Subject Category: 08
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    Report Date: June 1994
    No. Pages: 20
    Funding Organization: WU 533-02-34
    Keywords:      Emergency control; F-15 airplane; Flight test; Hydraulic failure; Propulsion-only control
    Notes: Presented as AIAA 94-2123 at the 7th Biennial Flight Test Conference, June 20-23, 1994, Colorado Springs, Colorado.


  14. AN APPROACH FOR EVALUATING THE DYNAMIC RESPONSE OF IN-FLIGHT THRUST CALIBRATION TECHNIQUES DURING THROTTLE TRANSIENTS , Technical Memorandum
    Authors: Ronald J. Ray
    Report Number: NASA-TM-4591
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: New fight test maneuvers and analysis techniques for evaluating the dynamic response of in-fight thrust models during throttle transient have been developed and validated. The approach is based on the aircraft and engine performance relationship between thrust and drag. Two fight test maneuvers, a throttle step and a throttle frequency sweep, were developed and used in the study. Graphical analysis techniques, including a frequency domain analysis method, were also developed and evaluated. They provide quantitative and qualitative results. Four thrust calculation methods were used to demonstrate and validate the test technique. Flight test applications on two high-performance aircraft confrmed the test methods as valid and accurate. These maneuvers and analysis techniques were easy to implement and use. Flight test results indicate the analysis techniques can identify the combined effects of model error and instrumentation response limitations on the calculated thrust value. The methods developed in this report provide an accurate approach for evaluating, validating, or comparing thrust calculation methods for dynamic fight applications.
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    Report Date: June 1994
    No. Pages: 27
    Funding Organization: WU 505-68
    Keywords:      Aircraft engine simulation; Aircraft engine testing; Aircraft flight test; Dynamic thrust; Performance; Thnrust; Thrust Validation; Unsteady Flow
    Notes: Also presented as AIAA 94-2115 at the 7th Biennial Flight Test Conference, June 20-23, 1994, Colorado Springs, Colorado.


  15. PRACTICAL APPLICATIONS OF CURRENT LOOP SIGNAL CONDITIONING , SPIE Conference Paper
    Authors: Karl F. Anderson
    Report Number: SPIE
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper describes a variety of practical application circuits based on the current loop signal conditioning paradigm. Equations defining the circuit response are also provided. The constant current loop is a fundamental signal conditioning circuit concept that can be implemented in a variety of configurations for resistance-based transducers, such as strain gages and resistance temperature devices. The circuit features signal conditioning outputs which are unaffected by extremely large variations in lead wire resistance, direct current frequency response, and inherent linearity with respect to resistance change. Sensitivity of this circuit is double that of a Wheatstone bridge circuit. Electrical output is zero for resistance change equals zero. The same excitation and output sense wires can serve multiple transducers. More application arrangements are possible with constant current loop signal conditioning than with the Wheatstone bridge.
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    Report Date: January 1994
    No. Pages: 20
    Funding Organization: n.a.
    Keywords:      Measurements; Resistance; Strain; Signal conditioning
    Notes: Also published at the NCSL 1994 Conference, July 31-Aug 4, 1994 in Chicago, IL and the ISA 40th International Symposium, May 1-5, 1994 in Baltimore, Maryland. and the Inst


  16. CONTROLLING FOREBODY ASYMMETRIES IN FLIGHT—EXPERIENCE WITH BOUNDARY LAYER TRANSITION STRIPS , Technical Memorandum
    Authors: David F. Fisher and Brent R. Cobleigh
    Report Number: NASA-TM-4595
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The NASA Dryden Flight Research Center has an ongoing program to investigate aircraft flight characteristics at high angles of attack. As part of this investigation, longitudinal boundary layer transition strips were installed on the F-18 HARV forebody, a preproduction F/A-18 radome with a nose-slice tendency, and the X-31 aircraft forebody and noseboom to reduce asymmetric yawing moments at high angles of attack. The transition strips were effective on the F-18 HARV at angles of attack above 60 deg. On the preproduction F/A-18 radome at an angle of attack near 50 deg the strips were not effective. When the transition strips were installed on the X-31 noseboom, a favorable effect was observed on the yawing moment dynamics but the magnitude of the yawing moments was not decreased.
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    Subject Category: 02
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    Report Date: July 1994
    No. Pages: 22
    Keywords:      Boundary layer separation; Boundary layer transition; F-18 aircraft; Flight tests; Forebodies; Pressure distribution; Vortices; X-31 aircraft; Yawing moments
    Notes: This was originally prepared as AIAA-94-1826 for the 6th Biennial Flight Test Conference, Colorado Springs, CO, June 20-23, 1994. WU 505-68-71


  17. DEVELOPMENT OF A LOW-ASPECT RATIO FIN FOR FLIGHT RESEARCH EXPERIMENTS , Technical Memorandum
    Authors: David M. Richwine and John H. Del Frate
    Report Number: NASA-TM-4596
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A second-generation flight test fixture, developed at NASA Dryden Flight Research Center, offers a generic testbed for aerodynamic and fluid mechanics research. The new fixture, a low-aspect ratio vertical fin shape mounted on the centerline of an F-15B aircraft lower fuselage, is designed for flight research at Mach numbers up to 2.0. The new fixture is a composite structure with a modular configuration and removable components for functional flexibility. This report describes the multidisciplinary design and analysis approach used to develop the fixture. The approach integrates conservative assumptions with simple analysis techniques to minimize the time and cost associated with its development. Presented are the principal disciplines required for this effort, which include aerodynamics, structures, stability, and operational considerations. In addition, preliminary results from the first phase of flight testing are presented. Acceptable directional stability and flow quality are documented and show agreement with predictions. Future envelope expansion activities will minimize current limitations so that the fixture can be used for a wide variety of high-speed aerodynamic and fluid mechanics research experiments.
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    Report Date: August 1994
    No. Pages: 15
    Funding Organization: WU 505-68
    Keywords:      Aerodynamic loads: F-15 directional stability; Flight test fixture; Flow visualization; Structural design; Structural ground testing; Water tunnel visualization studies
    Notes: This was originally prepared as AIAA-94-2133 for the 6th Biennial Flight Test Conference, Colorado Springs, CO, June 20-23, 1994.


  18. TRANSONIC FLIGHT TEST OF A LAMINAR FLOW LEADING EDGE WITH SURFACE EXCRESCENCES , Technical Memorandum
    Authors: Fanny A. Zuniga, Aaron Drake, Robert A. Kennelly, Jr., Dennis J. Koga and Russell V. Westphal
    Report Number: NASA-TM-4597
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A flight experiment, conducted at NASA Dryden Flight Research Center, investigated the effects of surface excrescences, specifically gaps and steps, on boundary-layer transition in the vicinity of a leading edge at transonic flight conditions. A natural laminar flow leading-edge model was designed for this experiment with a spanwise slot manufactured into the leading-edge model to simulate gaps and steps like those present at skin joints of small transonic aircraft wings. The leading-edge model was flown with the flight test fixture, a low-aspect ratio fin mounted beneath an F-104G aircraft. Test points were obtained over a unit Reynolds number range of 1.5 to 2.5 million/ft and a Mach number range of 0.5 to 0.8. Results for a smooth surface showed that laminar flow extended to approximately 12 in. behind the leading edge at Mach number 0.7 over a unit Reynolds number range of 1.5 to 2.0 million/ft. The maximum size of the gap-and-step configuration over which laminar flow was maintained consisted of two 0.06-in. gaps with a 0.02-in. step at a unit Reynolds number of 1.5 million/ft.
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    Report Date: August 1994
    No. Pages: 14
    Keywords:      Boundary-layer transition; Hot films; Laminar flow; Stanton gauge measurements; Surface excrescences
    Notes: Also presented as AIAA 94-2142 at the 6th Biennial Flight Test Conference, Colorado Springs, Colorado, June 20-23, 1994.


  19. X-29 FLIGHT CONTROL SYSTEM: LESSONS LEARNED , Technical Memorandum
    Authors: Robert Clarke, John J. Burken, John T. Bosworth and Jeffrey E. Bauer
    Report Number: NASA-TM-4598
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Two X-29A aircraft were flown at the NASA Dryden Flight Research Center over a period of eight years. The airplanes' unique features are the forward-swept wing, variable incidence close-coupled canard and highly relaxed longitudinal static stability (up to 35-percent negative static margin at subsonic conditions). This paper describes the primary flight control system and significant modifications made to this system, flight test techniques used during envelope expansion, and results for the low- and high-angle-of-attack programs. Throughout the paper, lessons learned will be discussed to illustrate the problems associated with the implementation of complex flight control systems.
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    Subject Category: 02
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    Report Date: June 1994
    No. Pages: 18
    Funding Organization: WU 505-64-30
    Keywords:      Flight control system; Flight-estimated stability margins; Static instability; X-29A airplane
    Notes: This was originally prepared for the AGARD Flight Mechanics Panel Symposium, Turin, Italy, May 9-12, 1994.


  20. WATER TUNNEL FLOW VISUALIZATION STUDY OF A 4.4 PERCENT SCALE X-31 FOREBODY , Technical Memorandum
    Authors: Brent Cobleigh and John Del Frate
    Report Number: NASA-TM-104276
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A water tunnel test of a 4.4 percent-scale, forebody-only model of the X-31 aircraft with different forebody strakes and nosebooms has been performed in the Flow Visualization Facility at the NASA Dryden Flight Research Center. The focus of the study was to determine the relative effects of the different configurations on the stability and symmetry of the high-angle-of-attack forebody vortex flow field. The clean, noseboom-off configuration resisted the development of asymmetries in the primary vortices through 70° angle of attack. The wake of the X-31 flight test noseboom configuration significantly degraded the steadiness of the primary vortex cores and promoted asymmetries. An alternate L-shaped noseboom mounted underneath the forebody had results similar to those seen with the noseboom-off configuration, enabling stable, symmetrical vortices up to 70° angle of attack. The addition of strakes near the radome tip along the waterline increased the primary vortex strength while it simultaneously caused the vortex breakdown location to move forward. Forebody strakes did not appear to significantly reduce the asymmetries in the forebody vortex field in the presence of the flight test noseboom.
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    Subject Category: 02
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    Report Date: September 1994
    No. Pages: 40
    Funding Organization: WU-533-02
    Keywords:      Asymmetry; Forebody; High angle of attack; Noseboom; Strake; Vortex flow; Water tunnel; X-31
    Notes: (Cobleigh: PRC Inc., Edwards, CA; Del Frate: NASA Dryden)


  21. DYNAMIC GROUND EFFECTS OF AN F-15 AIRCRAFT , Technical Memorandum
    Authors: Stephen Corda, Mark T. Stephenson, Frank W. Burcham, Jr. and Robert E. Curry
    Report Number: NASA-TM-4604
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests to determine the changes in the aerodynamic characteristics of an F-15 aircraft caused by dynamic ground effects are described. Data were obtained for low- and high-sink rates between 0.7 and 6.5 ft/sec and at two landing approach speeds and flap settings: 150 kn with the flaps down and 170 kn with the flaps up. Simple correlation curves are given for the change in aerodynamic coefficients because of ground effects as a function of sink rate. Ground effects generally caused an increase in the lift, drag, and nose-down pitching moment coefficients. The change in the lift coefficient increased from approximately 0.05 at the high-sink rate to approximately 0.10 at the low-sink rate. The change in the drag coefficient increased from approximately 0 to 0.03 over this decreasing sink rate range. No significant difference because of the approach configuration was evident for lift and drag; however, a significant difference in pitching moment was observed for the two approach speeds and flap settings. For the 170 kn with the flaps up configuration, the change in the nose-down pitching moment increased from approximately -0.008 to -0.016. For the 150 kn with the flaps down configuration, the change was from approximately -0.008 to -0.038.
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    Subject Category: 02
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    Report Date: September 1994
    No. Pages: 25
    Funding Organization: WU-533-02-31
    Keywords:      Aircraft aerodynamics; Aircraft approach and landing; F-15 aircraft; Flight testing; Ground effect; Propulsion-controlled aircraft
    Notes: Stephen Corda, PRC Inc., Edwards, California; Mark T. Stephenson, Frank W. Burcham, and Robert E. Curry, NASA Dryden Flight Research Center, Edwards, California.


  22. ON THE USE OF CONTROLS FOR SUBSONIC TRANSPORT PERFORMANCE IMPROVEMENT: OVERVIEW AND FUTURE DIRECTIONS , Technical Memorandum
    Authors: Glenn B. Gilyard and Martin Espana
    Report Number: NASA-TM-4605
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Increasing competition among airline manufacturers and operators has highlighted the issue of aircraft efficiency. Fewer aircraft orders have led to an all-out efficiency improvement effort among the manufacturers to maintain if not increase their share of the shrinking number of aircraft sales. Aircraft efficiency is important in airline profitability and is key if fuel prices increase from their current low. In a continuing effort to improve aircraft efficiency and develop an optimal performance technology base, NASA Dryden Flight Research Center developed and flight tested an adaptive performance seeking control system to optimize the quasi-steady-state performance of the F-15 aircraft. The demonstrated technology is equally applicable to transport aircraft although with less improvement. NASA Dryden, in transitioning this technology to transport aircraft, is specifically exploring the feasibility of applying adaptive optimal control techniques to performance optimization of redundant control effectors. A simulation evaluation of a preliminary control law optimizes wing-aileron camber for minimum net aircraft drag. Two submodes are evaluated: one to minimize fuel and the other to maximize velocity. This paper covers the status of performance optimization of the current fleet of subsonic transports; available integrated controls technologies are reviewed to define approaches using active controls. A candidate control law for adaptive performance optimization is presented along with examples of algorithm operation.
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    Subject Category: 02, 03
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    Report Date: August 1994
    No. Pages: 18
    Funding Organization: WU 505-69-10
    Keywords:      Aircraft performance; Cambered wings; Drag reduction; Flight optimization; Fuel consumption; Optimal control; Optimization
    Notes: This originally appeared as AIAA-94-3515 for the Atmospheric Flight Mechanics Conference,Scottsdale, Arizona, August 1-3, 1994.


  23. MEASUREMENT UNCERTAINTY AND FEASIBILITY STUDY OF A FLUSH AIRDATA SYSTEM FOR A HYPERSONIC FLIGHT EXPERIMENT , Technical Memorandum
    Authors: Stephen A. Whitmore and Timothy R. Moes
    Report Number: NASA-TM-4627
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Presented is a feasibility and error analysis for a hypersonic flush airdata system on a hypersonic flight experiment (HYFLITE). HYFLITE heating loads make intrusive airdata measurement impractical. Although this analysis is specifically for the HYFLITE vehicle and trajectory, the problems analyzed are generally applicable to hypersonic vehicles. A layout of the flush-port matrix is shown. Surface pressures are related airdata parameters using a simple aerodynamic model. The model is linearized using small perturbations and inverted using nonlinear least-squares. Effects of various error sources on the overall uncertainty are evaluated using an error simulation. Error sources modeled include boundary-layer/viscous interactions, pneumatic lag, thermal transpiration in the sensor pressure tubing, misalignment in the matrix layout, thermal warping of the vehicle nose, sampling resolution, and transducer error. Using simulated pressure data for input to the estimation algorithm, effects caused by various error sources are analyzed by comparing estimator outputs with the original trajectory. To obtain ensemble averages the simulation is run repeatedly and output statistics are compiled. Output errors resulting from the various error sources are presented as a function of Mach number. Final uncertainties with all modeled error sources included are presented as a function of Mach number.
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    Subject Category: 02
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    Report Date: June 1994
    No. Pages: 18
    Funding Organization: WU 505-68-40
    Keywords:      Airdata; Hypersonic flight; Least-squares; NASP; Parameter estimation
    Notes: This memorandum originally appeared as AIAA-94-1930 and was prepared for the Applied Aerodynamics Conference, Colorado Springs, CO, June 20-23, 1994.


  24. EFFECTS OF MASS ON AIRCRAFT SIDEARM CONTROLLER CHARACTERISTICS , Technical Memorandum
    Authors: Charles A. Wagner
    Report Number: NASA-TM-104277
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: When designing a flight simulator, providing a set of low mass variable-characteristic pilot controls can be very difficult. Thus, a strong incentive exists to identify the highest possible mass that will not degrade the validity of a simulation. The NASA Dryden Flight Research Center has conducted a brief flight program to determine the maximum acceptable mass (system inertia) of an aircraft sidearm controller as a function of force gradient. This information is useful for control system design in aircraft as well as development of suitable flight simulator controls. A modified Learjet with a variable-characteristic sidearm controller was used to obtain data. A boundary was defined between mass considered acceptable and mass considered unacceptable to the pilot. This boundary is defined as a function of force gradient over a range of natural frequencies. This investigation is limited to a study of mass-frequency characteristics only. Results of this investigation are presented in this paper.
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    Subject Category: 08
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    Report Date: September 1994
    No. Pages: 11
    Keywords:      Control dynamics; Control mass; Force gradient; Handling qualities; Natural frequency; Sidearm controllers
    Notes: N.A. WU 505-68-20


  25. HIGH-ANGLE-OF-ATTACK YAWING MOMENT ASYMMETRY OF THE X-31 AIRCRAFT FROM FLIGHT TEST , Contractor Report
    Authors: Brent R. Cobleigh
    Report Number: NASA-CR-186030
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Significant yawing moment asymmetries were encountered during the high-angle-of-attack envelope expansion of the two X-31 aircraft. These asymmetries led to position saturations of the thrust vector vanes and trailing-edge flaps during some of the dynamic stability axis rolling maneuvers at high angles of attack. This slowed the high-angle-of-attack envelope expansion and resulted in maneuver restrictions. Several aerodynamic modifications were made to the X-31 forebody with the goal of minimizing the asymmetry. A method for determining the yawing moment asymmetry from flight data was developed and an analysis of the various configuration changes completed. The baseline aircraft were found to have significant asymmetries above 45 deg-angle of attack with the largest asymmetry typically occurring around 60 deg-angle of attack. Applying symmetrical boundary-layer transition strips along the forebody sides increased the magnitude of the asymmetry and widened the angle-of-attack range over which the largest asymmetry acted. Installing longitudinal forebody strakes and rounding the sharp nose of the aircraft caused the yawing moment asymmetry magnitude to be reduced. The transition strips and strakes made the asymmetry characteristic of the aircraft more repeatable than the clean forebody configuration. Although no geometric differences between the aircraft were known, ship 2 consistently had larger yawing moment asymmetries than ship 1.
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    Subject Category: 05
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    Report Date: September 1994
    No. Pages: 33
    Funding Organization: RTOP 533-02
    Keywords:      Asymmetries; Forebody; High angle of attack; Noseboom; Reynolds number; Strakes; Tactical utility; Transition strips; X-31; Yawing moment
    Notes: Presented as AIAA 94-1803 at the Applied Aerodynamics Conference, Colorado Springs, CO, June 20-23, 1994.


  26. SHEAR BUCKLING ANALYSIS OF A HAT-STIFFENED PANEL , Technical Memorandum
    Authors: William L. Ko and Raymond H. Jackson
    Report Number: NASA-TM-4644
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A buckling analysis was performed on a hat-stiffened panel subjected to shear loading. Both local buckling and global buckling were analyzed. The global shear buckling load was found to be several times higher than the local shear buckling load. The classical shear buckling theory for a flat plate was found to be useful in predicting the local shear buckling load of the hat-stiffened panel, and the predicted local shear buckling loads thus obtained compare favorably with the results of finite element analysis.
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    Subject Category: 39
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    Report Date: November 1994
    No. Pages: 19
    Funding Organization: WU 505-63-40
    Keywords:      Hat-stiffened panel; Global buckling; Local buckling; Minimum energy method; Shear buckling
    Notes: n.a.


  27. DETERMINATION OF STORES POINTING ERROR DUE TO WING FLEXIBILITY UNDER FLIGHT LOAD , Technical Memorandum
    Authors: William A. Lokos, Catherine M. Bahm and Robert A. Heinle
    Report Number: NASA-TM-4646
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The in-flight elastic wing twist of a fighter-type aircraft was studied to provide for an improved on-board realtime computed prediction of pointing variations of three wing store stations. This is an important capability to correct sensor pod alignment variation or to establish initial conditions of iron bombs or smart weapons prior to release. The original algorithm was based upon coarse measurements. The electro-optical Flight Deflection Measurement System measured the deformed wing shape in flight under maneuver loads to provide a higher resolution database from which an improved twist prediction algorithm could be developed. The FDMS produced excellent repeatable data. In addition, a NASTRAN finite-element analysis was performed to provide additional elastic deformation data. The FDMS data combined with the NASTRAN analysis indicated that an improved prediction algorithm could be derived by using a different set of aircraft parameters, namely normal acceleration, stores configuration, Mach number, and gross weight.
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    Subject Category: 05
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    Report Date: December 1994
    No. Pages: 24
    Funding Organization: WU 533-02-31
    Keywords:      Elastic wing twist; F-16 wing deformations; In-flight deflection measurement; Wing store alignment
    Notes: Presented as AIAA-94-2112 at the 7th Biennial Flight Test Conference, Colorado Springs, CO, June 20-23, 1994.


  28. IN-FLIGHT IMAGING OF TRANSVERSE GAS JETS INJECTED INTO TRANSONIC AND SUPERSONIC CROSSFLOWS: DESIGN AND DEVELOPMENT
    Authors: Kon-Sheng Charles Wang
    Report Number: NASA-CR-186031
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The design and development of an airborne flight-test experiment to study nonreacting gas jets injected transversely into transonic and supersonic crossflows is presented. Free-stream/crossflow Mach numbers range from 0.8 to 2.0. Planar laser-induced fluorescence (PLIF) of an iodine-seeded nitrogen jet is used to visualize the jet flow. Time-dependent images are obtained with a high-speed intensified video camera synchronized to the laser pulse rate. The entire experimental assembly is configured compactly inside a unique flight-test-fixture (FTF) mounted under the fuselage of the F-104G research aircraft, which serves as a 'flying wind tunnel' at NASA Dryden Flight Research Center. The aircraft is flown at predetermined speeds and altitudes to permit a perfectly expanded (or slightly underexpanded) gas jet to form just outside the FTF at each free stream Mach number. Recorded gas jet images are then digitized to allow analysis of jet trajectory, spreading, and mixing characteristics. Comparisons will be made with analytical and numerical predictions. (Results presented in AIAA CP-95-0516). This study shows the viability of applying highly sophisticated ground-based flow diagnostic techniques to flight-test vehicle platforms that can achieve a wide range of thermo/fluid dynamic conditions. Realistic flow environments, high enthalpies, unconstrained flow-fields, and moderate operating costs are also realized, in contrast to traditional wind-tunnel testing.
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    Subject Category: 02, 34
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    Report Date: November 1994
    No. Pages: 85
    Funding Organization: WU 505-68-53, NCC 2-374
    Keywords:      F-104; Flight test; Flight test fixture; Flow visualization; Imaging; Injection; Nd: YAG laser; Planar laser-induced fluorescence (PLIF); Supersonic; Transonic; Transverse jet
    Notes: Tech monitors Dr. Iliff and Bowers. This report was originally prepared as a project submitted in partial satisfaction of the requirements for the degree Master of Science in Mechanical Engineering at University of Calif., Los Angeles, March 1993.


  29. NUMERICAL MODELING OF A CRYOGENIC FLUID WITHIN A FUEL TANK , Technical Memorandum
    Authors: Donald S. Greer
    Report Number: NASA-TM-4651
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The computational method developed to study the cryogenic fluid characteristics inside a fuel tank in a hypersonic aircraft is presented. The model simulates a rapid draining of the tank by modeling the ullage vapor and the cryogenic liquid with a moving interface. A mathematical transformation was developed and applied to the Navier-Stokes equations to account for the moving interface. The formulation of the numerical method is a transient hybrid explicit-implicit technique where the pressure term in the momentum equations is approximated to first order in time by combining the continuity equation with an ideal equation of state.
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    Subject Category: 34
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    Report Date: October 1994
    No. Pages: 15
    Funding Organization: WU 505-70-63
    Keywords:      Compressible flow; Computation fluid dynamics (CFD); Fluid flow; Fluid mechanics; Gas dynamics; Numerical modeling; Unsteady flow
    Notes: Presented at the Second Thermal Structures Conference, Charlottesville, Virginia, October 18-21, 1994.


  30. DEVELOPMENT OF A MULTICOMPONENT FORCE AND MOMENT BALANCE FOR WATER TUNNEL APPLICATIONS, VOLUME I , Contractor Report
    Authors: Carlos J. Suarez, Gerald N. Malcolm, Brian R. Kramer, Brooke C. Smith and Bert F. Ayers
    Report Number: NASA-CR-4642
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The principal objective of this research effort was to develop a multicomponent strain gauge balance to measure forces and moments on models tested in flow visualization water tunnels. An internal balance was designed that allows measuring normal and side forces, and pitching, yawing and rolling moments (no axial force). The five-components to applied loads, low interactions between the sections and no hysteresis. Static experiments (which are discussed in this Volume) were conducted in the Eidetics water tunnel with delta wings and a model of the F/A-18. Experiments with the F/A-18 model included a thorough baseline study and investigations of the effect of control surface deflections and of several Forebody Vortex Control (FVC) techniques. Results were compared to wind tunnel data and, in general, the agreement is very satisfactory. The results of the static tests provider confidence that loads can be measured accurately in the water tunnel with a relatively simple multi-component internal balance. Dynamic experiments were also performed using the balance, and the results are discussed in detail in Volume II of this report.
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    Subject Category: 34
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    Report Date: December 1994
    No. Pages: 100
    Funding Organization: WU 505-5-953, NAS2-13571
    Keywords:      Static and dynamic experiments; Water tunnel force/moment balance
    Notes: NASA technical monitor John Del Frate


  31. EULERIAN-LAGRANGIAN SIMULATIONS OF TRANSONIC FLUTTER INSTABILITIES , Conference Report
    Authors: Oddvar O. Bendiksen
    Report Number: H-2293
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: This paper presents an overview of recent applications of Eulerian-Lagrangian computational schemes in simulating transonic flutter instabilities. In this approach, the fluid-structure system is treated as a single continuum dynamics problem, by switching from an Eulerian to Lagrangian formulation at the fluid-structure boundary. This computational approach effectively eliminates the phase integration errors associated with previous methods, where the fluid and structure are integrated sequentially using different schemes. The formulation is based on Hamilton's Principle in mixed coordinates, and both finite volume and finite element discretization schemes are considered. Results from numerical simulations of transonic flutter instabilities are presented for isolated wings, thin panels, and turbomachinery blades. The results suggest that the method is capable of reproducing the energy exchange between the fluid and the structure with significantly less error than existing methods. Localized flutter modes and panel flutter modes involving traveling waves can also be simulated effectively with no a priori knowledge of the type of instability involved.
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    Report Date: November 1994
    No. Pages: 35
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: At Symp. on Aeroelasticity and Fluid-Structure Interaction Problems, ASME, Chicago, IL, 11/6/94. Doc. result of work sponsored by NASA Dryden, performed through UCLA, Flight Sys. Res. Ctr. Contract monitor, R. Quinn; funding facilitated by K. Iliff.


  32. FRICTION FACTOR FOR FLOW IN RECTANGULAR DUCTS WITH ONE SIDE RIB-ROUGHENED , Journal Article
    Authors: B. Youn, C. Yuen and A. F. Mills
    Report Number: H-2294
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Numerical simulations of incompressible turbulent flow through rectangular ducts with one side rib-roughened were performed to determine pressure drop. The 'PHOENICS' software package was used for the computations, which required provision of a wall function for transverse rib-roughened surfaces. The present study was conducted in the range of 10(5)(superscript) =/< Reynolds number =/< 10(7)(superscript), 0.01 =/< rib height to hydraulic diameter ratio =/< 0.04, 10 =/< pitch to rib height ratio =/< 40. Using the numerical results, friction factor charts for various aspect ratios were generated. The numerical results agreed well with experimental data that was obtained for 10(5)(superscript) < Reynolds number < 2 x 10(5)(superscript). In addition, a scheme for predicting friction factor using existing correlations for smooth and rough walls was developed.
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    Report Date: September 1994
    No. Pages: 6
    Funding Organization: NASA Grant number NCC-2-374
    Keywords:      n.a.
    Notes: Journal of Fluids Engineering, Vol. 116, 1994, pp. 488-493. 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.