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  1. EVALUATION OF VARIOUS THRUST CALCULATION TECHNIQUES ON AN F404 ENGINE , Technical Paper
    Authors: Ronald J. Ray
    Report Number: NASA-TP-3001
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In support of performance testing of the X-29A aircraft at the NASA-Ames, various thrust calculation techniques were developed and evaluated for use on the F404-GE-400 engine. The engine was thrust calibrated at NASA-Lewis. Results from these tests were used to correct the manufacturer's in-flight thrust program to more accurately calculate thrust for the specific test engine. Data from these tests were also used to develop an independent, simplified thrust calculation technique for real-time thrust calculation. Comparisons were also made to thrust values predicted by the engine specification model. Results indicate uninstalled gross thrust accuracies on the order of 1 to 4 percent for the various in-flight thrust methods. The various thrust calculations are described and their usage, uncertainty, and measured accuracies are explained. In addition, the advantages of a real-time thrust algorithm for flight test use and the importance of an accurate thrust calculation to the aircraft performance analysis are described. Finally, actual data obtained from flight test are presented.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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    Report Date: April 1990
    No. Pages: 31
    Keywords:      Calibrating; Engine tests; Flight tests; Performance prediction; Real time operation; Thrust


  2. THE EFFECTIVENESS OF VANE-AILERON EXCITATION IN THE EXPERIMENTAL DETERMINATION OF FLUTTER SPEED BY PARAMETER IDENTIFICATION , Technical Paper
    Authors: E. Nissim
    Report Number: NASA-TP-2971
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The effectiveness of aerodynamic excitation is evaluated analytically in conjunction with the experimental determination of flutter dynamic pressure by parameter identification. Existing control surfaces were used, with an additional vane located at the wingtip. The equations leading to the identification of the equations of motion were reformulated to accommodate excitation forces of aerodynamic origin. The aerodynamic coefficients of the excitation forces do not need to be known since they are determined by the identification procedure. The 12 degree-of-freedom numerical example treated in this work revealed the best wingtip vane locations, and demonstrated the effectiveness of the aileron-vane excitation system. Results from simulated data gathered at much lower dynamic pressures (approximately half the value of flutter dynamic pressure) predicted flutter dynamic pressures with 2-percent errors.
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    Subject Category: A02
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    Report Date: January 1990
    No. Pages: 24


  3. OPTIMUM ELEMENT DENSITY STUDIES FOR FINITE-ELEMENT THERMAL ANALYSIS OF HYPERSONICAIRCRAFT STRUCTURES , Technical Memorandum
    Authors: William L. Ko, Timothy Olona and Kyle M. Muramoto
    Report Number: NASA-TM-4163
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Different finite element models previously set up for thermal analysis of the space shuttle orbiter structure are discussed and their shortcomings identified. Element density criteria are established for the finite element thermal modelings of space shuttle orbiter-type large, hypersonic aircraft structures. These criteria are based on rigorous studies on solution accuracies using different finite element models having different element densities set up for one cell of the orbiter wing. Also, a method for optimization of the transient thermal analysis computer central processing unit (CPU) time is discussed. Based on the newly established element density criteria, the orbiter wing midspan segment was modeled for the examination of thermal analysis solution accuracies and the extent of computation CPU time requirements. The results showed that the distributions of the structural temperatures and the thermal stresses obtained from this wing segment model were satisfactory and the computation CPU time was at the acceptable level. The studies offered the hope that modeling the large, hypersonic aircraft structures using high-density elements for transient thermal analysis is possible if a CPU optimization technique was used.
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    Subject Category: 34
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    Report Date: January 1990
    No. Pages: 36
    Keywords:      Aircraft structures; Finite element method; Hypersonic aircraft; Mathematical models; Thermal analysis; Thermal stresses


  4. OUTPUT MODEL-FOLLOWING CONTROL SYNTHESIS FOR AN OBLIQUE-WING AIRCRAFT , Technical Memorandum
    Authors: Joseph W. Pahle
    Report Number: NASA-TM-100454
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Recent interest in oblique-wing aircraft has focused on the potential aerodynamic performance advantage of a variable-skew oblique wing over a conventional or symmetric sweep wing. Unfortunately, the resulting asymmetric configuration has significant aerodynamic and inertial cross-coupling between the aircraft longitudinal and lateral-directional axes. Presented here is a decoupling control law synthesis technique that integrates stability augmentation, decoupling, and the direct incorporation of desired handling qualities into an output feedback controller. The proposed design technique uses linear quadratic regulator concepts in the framework of explicit model following. The output feedback strategy used is a suboptimal projection from the state space to the output space. Dynamics are then introduced into the controller to improve steady-state performance and increase system robustness. Closed-loop performance is shown by application of the control laws to the linearized equations of motion and nonlinear simulation of an oblique-wing aircraft.
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    Subject Category: 08
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    Report Date: April 1990
    No. Pages: 30
    Keywords:      Aerodynamic characteristics; Control theory; Controllers; Feedback control; Oblique wings; Optimal control; Stability augmentation; Swept wings


  5. EFFECTS OF WING SWEEP ON BOUNDARY-LAYER TRANSITIONFOR A SMOOTH F-14A WING AT MACH NUMBERS FROM 0.700 TO0.825 , Technical Memorandum
    Authors: Bianca Trujillo Anderson and Robert R. Meyer, Jr.
    Report Number: NASA-TM-101712
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The results are discussed of the variable sweep transition flight experiment (VSTFE). The VSTFE was a natural laminar flow experiment flown on the swing wing F-14A aircraft. The main objective of the VSTFE was to determine the effects of wing sweep on boundary layer transition at conditions representative of transport aircraft. The experiment included the flight testing of two laminar flow wing gloves. Glove 1 was a cleanup of the existing F-14A wing. Glove 2, not discussed herein, was designed to provide favorable pressure distributions for natural laminar flow at Mach number (M) 0.700. The transition locations presented for glove 1 were determined primarily by using hot film sensors. Boundary layer rake data was provided as a supplement. Transition data were obtained for leading edge wing sweeps of 15, 20, 25, 30, and 35 degs, with Mach numbers ranging from 0.700 to 0.825, and altitudes ranging from 10,000 to 35,000 ft. Results show that a substantial amount of laminar flow was maintained at all the wing sweeps evaluated. The maximum transition Reynolds number of 13.7 x 10(exp 6) was obtained for the condition of 15 deg of sweep, M = 0.800, and an altitude of 20,000 ft.
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    Subject Category: 34
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    Report Date: May 1990
    No. Pages: 48
    Keywords:      Boundary layer transition; Flow distribution; Laminar flow; Mach number; Variable sweep wings


  6. EXPERIMENTAL CHARACTERIZATION OF THE EFFECTS OF PNEUMATIC TUBING ON UNSTEADY PRESSUREMEASUREMENTS , Technical Memorandum
    Authors: Stephen A. Whitmore, William T. Lindsey, Robert E. Curry and Glenn B. Gilyard
    Report Number: NASA-TM-4171
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Advances in aircraft control system designs have, with increasing frequency, required that air data be used as flight control feedback. This condition requires that these data be measured with accuracy and high fidelity. Most air data information is provided by pneumatic pressure measuring sensors. Typically unsteady pressure data provided by pneumatic sensing systems are distorted at high frequencies. The distortion is a result of the pressure being transmitted to the pressure sensor through a length of connective tubing. The pressure is distorted by frictional damping and wave reflection. As a result, air data provided all-flush, pneumatically sensed air data systems may not meet the frequency response requirements necessary for flight control augmentation. Both lab and flight test were performed at NASA-Ames to investigate the effects of this high frequency distortion in remotely located pressure measurement systems. Good qualitative agreement between lab and flight data are demonstrated. Results from these tests are used to describe the effects of pneumatic distortion in terms of a simple parametric model.
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    Subject Category: 06
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    Report Date: March 1990
    No. Pages: 31
    Keywords:      Air data systems; Control systems design; Pipes (tubes); Penumatic probes; Pressure measurement; Signal distortion; Unsteady flow


  7. WIND-TUNNEL INVESTIGATION OF A FLUSH AIRDATA SYSTEM AT MACH NUMBERS FROM 0.7 TO 1.4 , Technical Memorandum
    Authors: Terry J. Larson, Timothy R. Moes and Paul M. Siemers, III
    Report Number: NASA-TM-101697
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flush pressure orifices installed on the nose section of a 1/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wing-tunnel tests were conducted in the 11- by 11-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which were published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2 deg increments from -4 deg to 20 deg. Sideslip angles were varied in 4 deg increments from -8 deg to 8 deg. Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.
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    Report Date: April 1990
    No. Pages: 35
    Keywords:      Air data systems; Algorithms; Flight tests; Flushing; Wind tunnel tests; Wind tunnels


  8. DESIGN OF CONTROL LAWS FOR FLUTTER SUPPRESSION BASED ON THE AERODYNAMIC ENERGY CONCEPT AND COMPARISONS WITH OTHER DESIGN METHODS , Technical Paper
    Authors: Eli Nissim (Technion - Israel Inst. of Tech.)
    Report Number: NASA-TP-3056
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The aerodynamic energy method is used to synthesize control laws for NASA's drone for aerodynamic and structural testing-aerodynamic research wing 1 (DAST-ARW1) mathematical model. The performance of these control laws in terms of closed-loop flutter dynamic pressure, control surface activity, and robustness is compared with other control laws that relate to the same model. A control law synthesis technique that makes use of the return difference singular values is developed. It is based on the aerodynamic energy approach and is shown to yield results that are superior to those results given in the literature and are based on optimal control theory. Nyquist plots are presented, together with a short discussion regarding the relative merits of the minimum singular value as a measure of robustness as compared with the more traditional measure involving phase and gain margins.
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    Subject Category: 39
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    Report Date: October 1990
    No. Pages: 59
    Keywords:      Aeroelastic research wings; Control systems design; Control theory; Energy methods; Flutter analysis; Vibration damping


  9. FLIGHT CHARACTERISTICS OF A MODIFIED SCHWEIZER SGS 1-36 SAILPLANE AT LOW AND VERY HIGH ANGLES OF ATTACK , Technical Paper
    Authors: Alex G. Sim
    Report Number: NASA-TP-3022
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A manned flight research program using a modified sailplane was conducted to very high angles of attack at the NASA-Ames. Piloting techniques were established that enabled the pilot to attain and stabilize on an angle of attack in the 30 to 72 deg range. Aerodynamic derivatives were estimated from the flight data for both low and very high angles of attack and are compared to wind tunnel data. In addition, limited performance and trim data are presented.
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    Subject Category: 08
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    Report Date: July 1990
    No. Pages: 48
    Keywords:      Aerodynamic stability; Angle of attack; Flight characteristics; Gliders; Parameter identification; Pilot performance


  10. EFFECTS OF WING SWEEP ON IN-FLIGHT BOUNDARY-LAYERTRANSITION FOR A LAMINAR FLOW WING AT MACH NUMBERS FROM 0.60TO 0.79 , Technical Memorandum
    Authors: Bianca Trujillo Anderson and Robert R. Meyer, Jr.
    Report Number: NASA-TM-101701
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The variable sweep transition flight experiment (VSTFE) was conducted on an F-14A variable sweep wing fighter to examine the effect of wing sweep on natural boundary layer transition. Nearly full span upper surface gloves, extending to 60 percent chord, were attached to the F-14 aircraft's wings. The results are presented of the glove 2 flight tests. Glove 2 had an airfoil shape designed for natural laminar flow at a wing sweep of 20 deg. Sample pressure distributions and transition locations are presented with the complete results tabulated in a database. Data were obtained at wing sweeps of 15, 20, 25, 30, and 35 deg, at Mach numbers ranging from 0.60 to 0.79, and at altitudes ranging from 10,000 to 35,000 ft. Results show that a substantial amount of laminar flow was maintained at all the wing sweeps evaluated. The maximum transition Reynolds number obtained was 18.6 x 10(exp 6) at 15 deg of wing sweep, Mach 0.75, and at an altitude of 10,000 ft.
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    Subject Category: 34
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    Report Date: July 1990
    No. Pages: 37
    Keywords:      Boundary layer transition; Flow distribution; Laminar flow; Mach number; Pressure distribution; Variable sweep wings


  11. OPEN-MODE DELAMINATION STRESS CONCENTRATIONS IN HORSESHOE AND ELLIPTIC COMPOSITE CURVED BARS SUBJECTED TO END FORCES , Technical Memorandum
    Authors: William L. Ko and Raymond H. Jackson
    Report Number: NASA-TM-4164
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The multilayer theory of anisotropic elasticity and a finite element method were used to analyze the open-mode delamination stress concentrations in horseshoe and elliptic laminated composite curved bars. Two types of laminations, solid laminations and sandwich laminations, were analyzed. It was found that the open-mode delamination stress concentration could be greatly increased in these two types of curved bars by decreasing their aspect ratios. The open-mode delamination stress concentration generated in the solid laminations was found to be far more severe than that generated in the sandwich laminations. The horseshoe curved bar may be used to determine both the open-mode delamination strength of solidly laminated composites and the open-mode debonding strength of sandwiched laminated composites. However, the elliptic curved bar is only good for determining the open-mode delamination strength of solidly laminated composites.
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    Report Date: January 1990
    No. Pages: 25
    Keywords:      Bars; Delaminating; Finite element method; Sandwich structures; Stress concentration


  12. WATER-TUNNEL STUDY RESULTS OF A TF/A-18 AND F/A-18 CANOPY FLOW VISUALIZATION , Technical Memorandum
    Authors: Steven A. Johnson and David F. Fisher
    Report Number: NASA-TM-101705
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A water tunnel study examining the influence of canopy shape on canopy and leading edge extension flow patterns was initiated. The F/A-18 single-place canopy model and the TF/A-18 two place canopy model were the study subjects. Plan view and side view photographs showing the flow patterns created by injected colored dye are presented for 0 deg and 5 deg sideslip angles. Photographs taken at angle of attack and sideslip conditions correspond to test departure points found in flight test. Flight experience has shown that the TF/A-18 airplane departs in regions where the F/A-18 airplane is departure-resistant. The study results provide insight into the differences in flow patterns which may influence the resulting aerodynamics of the TF/A-18 and F/A-18 aircraft. It was found that at 0 deg sideslip, the TF/A-18 model has more downward flow on the sides of the canopy than the F/A-18 model. This could be indicative of flow from the leading edge extension (LEX) vortexes impinging on the sides of the wider TF/A-18 canopy. In addition, the TF/A-18 model has larger areas of asymmetric separated and unsteady flow on the LEXs and fuselage, possibly indicating a lateral and directional destabilizing effect at the conditions studied.
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    Report Date: March 1990
    No. Pages: 29
    Keywords:      Aerodynamics; Canopies; F-18 aircraft; Flow distribution; Flow visualization; Fuselages


  13. IN-FLIGHT INVESTIGATION OF SHUTTLE TILE PRESSURE ORIFICE INSTALLATIONS , Technical Memorandum
    Authors: Timothy R. Moes and Robert R. Meyer, Jr.
    Report Number: NASA-TM-4219
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: To determine shuttle orbiter wing loads during ascent, wing load instrumentation was added to Columbia (OV-102). This instrumentation included strain gages and pressure orifices on the wing. The loads derived from wing pressure measurements taken during STS 61-C did not agree with those derived from strain gage measurements or with the loads predicted from the aerodynamic database. Anomalies in the surface immediately surrounding the pressure orifices in the thermal protection system (TPS) tiles were one possible cause of errors in the loads derived from wing pressure measurements. These surface anomalies were caused by a ceramic filler material which was installed around the pressure tubing. The filler material allowed slight movement of the TPS tile and pressure tube as the airframe flexed and bent under aerodynamic loads during ascent and descent. Postflight inspection revealed that this filler material had protruded from or receeded beneath the surface, causing the orifice to lose its flushness. Flight tests were conducted at NASA Ames Research Center Dryden Flight Research Facility to determine the effects of any anomaly in surface flushness of the orifice installation on the measured pressures at Mach numbers between 0.6 and 1.4. An F-104 aircraft with a flight test fixture mounted beneath the fuselage was used for these flights. Surface flushness anomalies typical of those on the orbiter after flight (STA 61-C) were tested. Also, cases with excessive protrusion and recession of the filler material were tested. This report shows that the anomalies in STS 61-C orifice installations adversely affected the pressure measurements. But the magnitude of the affect was not great enough to account for the discrepancies with the strain gage measurements and the aerodynamic predictions.
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    Subject Category: 34
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    Report Date: September 1990
    No. Pages: 38
    Keywords:      Ceramics; In-flight monitoring; Installing; Orifices; Space shuttle orbiters; Tiles; Wing loading; Wings


  14. IN-FLIGHT FLOW VISUALIZATION CHARACTERISTICS OF THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE AT HIGHANGLES OF ATTACK , Technical Memorandum
    Authors: David F. Fisher, John H. Del Frate and David M. Richwine
    Report Number: NASA-TM-4193
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Surface and off-surface flow visualization techniques were used to visualize the 3-D separated flows on the NASA F-18 high alpha research vehicle at high angles of attack. Results near the alpha = 25 to 26 deg and alpha = 45 to 49 deg are presented. Both the forebody and leading edge extension (LEX) vortex cores and breakdown locations were visualized using smoke. Forebody and LEX vortex separation lines on the surface were defined using an emitted fluid technique. A laminar separation bubble was also detected on the nose cone using the emitted fluid technique and was similar to that observed in the wind tunnel test, but not as extensive. Regions of attached, separated, and vortical flow were noted on the wing and the leading edge flap using tufts and flow cones, and compared well with limited wind tunnel results.
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    Subject Category: 02
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    Report Date: May 1990
    No. Pages: 34
    Keywords:      F-18; Flight test; Flow visualization; Separated flow; Separation; Transition; Vortex; Vortex core; Wing rock


  15. AIRDATA CALIBRATION OF A HIGH-PERFORMANCE AIRCRAFT FOR MEASURING ATMOSPHERIC WINDPROFILES , Technical Memorandum
    Authors: Edward A. Haering, Jr.
    Report Number: NASA-TM-101714
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The research airdata system of an instrumented F-104 aircraft has been calibrated to measure winds aloft in support of the space shuttle wind measurement investigation at the National Aeronautics and Space Administration Ames Research Center Dryden Flight Research Facility. For this investigation, wind measurement accuracies comparable to those obtained from Jimsphere balloons were desired. This required an airdata calibration more accurate than needed for most aircraft research programs. The F-104 aircraft was equipped with a research pilot-static noseboom with integral angle-of-attack and flank angle-of-attack vanes and a ring-laser-gyro inertial reference unit. Tower fly-bys and radar acceleration-decelerations were used to calibrate Mach number and total temperature. Angle of attack and angle of sideslip were calibrated with a trajectory reconstruction technique using a multiple-state linear Kalman filter. The F-104 aircraft and instrumentation configuration, flight test maneuvers, data corrections, calibration techniques, and resulting calibrations and data repeatability are presented. Recommendations for future airdata systems on aircraft used to measure winds aloft are also given.
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    Subject Category: 05
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    Report Date: January 1990
    No. Pages: 27
    Notes: AIAA Paper 90-0230. Presented at the AIAA 28th Aerospace Sciences Meeting, Reno, Nevada, January 8-11, 1990.


  16. PRELIMINARY RESULTS FROM A SUBSONIC HIGH ANGLE-OF-ATTACK FLUSH AIRDATA SENSING (HI-FADS)SYSTEM: DESIGN, CALIBRATION, AND FLIGHT TEST EVALUATION , Technical Memorandum
    Authors: Stephen A. Whitmore, Timothy R. Moes and Terry J. Larson
    Report Number: NASA-TM-101713
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A nonintrusive high angle-of-attack flush airdata sensing (HI-FADS) system was installed and flight-tested on the F-18 high alpha research flight vehicle. The system is a matrix of 25 pressure orifices in concentric circles on the nose of the vehicle. The orifices determine angles of attack and sideslip, Mach number, and pressure altitude. Pressure was transmitted from the orifices to an electronically scanned pressure module by lines of pneumatic tubing. The HI-FADS system was calibrated and demonstrated using dutch roll flight maneuvers covering large Mach, angle-of-attack, and sideslip ranges. Reference airdata for system calibration were generated by a minimum variance estimation technique blending measurements from two wingtip airdata booms with inertial velocities, aircraft angular rates and attitudes, precision radar tracking, and meteorological analyses. The pressure orifice calibration was based on identifying empirical adjustments to modified Newtonian flow on a hemisphere. Calibration results are presented. Flight test results used all 25 orifices or used a subset of 9 orifices. Under moderate maneuvering conditions, the HI-FADS system gave excellent results over the entire subsonic Mach number range up to 55 deg angle of attack. The internal pneumatic frequency response of the system is accurate to beyond 10 Hz. Aerodynamic lags in the aircraft flow field caused some performance degradation during heavy maneuvering.
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    Subject Category: 06
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    Report Date: January 1990
    No. Pages: 13
    Keywords:      Air data systems; Aircraft design; Aircraft maneuvers; Angle of attack; Flight tests; Subsonic aircraft
    Notes: AIAA Paper 90-0232. Presented at the 28th AIAA Aerospace Sciences Meeting, Reno, Nevada, January 8-11, 1990.


  17. COMPENSATING FOR PNEUMATIC DISTORTION IN PRESSURE SENSING DEVICES , Technical Memorandum
    Authors: Stephen A. Whitmore and Cornelius T. Leondes
    Report Number: NASA-TM-101716
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A technique of compensating for pneumatic distortion in pressure sensing devices was developed and verified. This compensation allows conventional pressure sensing technology to obtain improved unsteady pressure measurements. Pressure distortion caused by frictional attenuation and pneumatic resonance within the sensing system makes obtaining unsteady pressure measurements by conventional sensors difficult. Most distortion occurs within the pneumatic tubing which transmits pressure impulses from the aircraft's surface to the measurement transducer. To avoid pneumatic distortion, experiment designers mount the pressure sensor at the surface of the aircraft, (called in-situ mounting). In-situ transducers cannot always fit in the available space and sometimes pneumatic tubing must be run from the aircraft's surface to the pressure transducer. A technique to measure unsteady pressure data using conventional pressure sensing technology was developed. A pneumatic distortion model is reduced to a low-order, state-variable model retaining most of the dynamic characteristics of the full model. The reduced-order model is coupled with results from minimum variance estimation theory to develop an algorithm to compensate for the effects of pneumatic distortion. Both postflight and real-time algorithms are developed and evaluated using simulated and flight data.
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    Subject Category: 05
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    Report Date: January 1990
    No. Pages: 21
    Keywords:      Dynamic characteristics; Flight control; Flow distortion; Pneumatic probes; Pressure measurement; Pressure sensors
    Notes: Presented at the AIAA 28th Aerospace Sciences Meeting, Reno, Nevada, January 8-11, 1990.


  18. AN IN-FLIGHT INTERACTION OF THE X-29A CANARD AND FLIGHT CONTROL SYSTEM , Technical Memorandum
    Authors: Michael W. Kehoe , Edward J. Laurie and Lisa J. Bjarke
    Report Number: NASA-TM-101718
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Many of today's high performance airplanes use high gain, digital flight control systems. These sytems are liable to couple with the aircraft's structural dynamics and aerodynamics to cause an aeroservoelastic interaction. These interactions can be stable or unstable depending upon damping and phase relationships within the system. The details of an aeroservoelastic interaction experienced in flight by the X- 29A forward-swept wing airplane. A 26.5-Hz canard pitch mode response was aliased by the digital sampling rate in the canard position feedback loop of the flight control system, resulting in a 13.5-Hz signal being commanded to the longitudinal control surfaces. The amplitude of this commanded signal increased as the wear of the canard seals increased, as the feedback path gains were increased, and as the canard aerodynamic loading decreased. The resultant control surface deflections were of sufficient amplitude to excite the structure. The flight data presented shows the effect of each component (structural dynamics, aerodynamics, and flight control system) for this aeroservoelastic interaction.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
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    Report Date: April 1990
    No. Pages: 21
    Keywords:      Aerodynamic loads; Canard configurations; Dynamic response; Dynamic structural analysis; Flight control; X-29 aircraft
    Notes: Presented at the AIAA Dynamics Specialists Meeting, Long Beach, CA, April 5-6, 1990.


  19. IN-FLIGHT FLOW FIELD ANALYSIS ON THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE WITHCOMPARISONS TO GROUND FACILITY DATA , Conference Paper
    Authors: John H. Del Frate and Fanny A. Zuniga
    Report Number: H-1592
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In-flight flow visualization results of the vortical flow on the forebody and leading-edge extensions (LEX) of an F-18 research aircraft have been presented for angles of attack from 15.8 to 42.5 deg and for sideslip angles up to 7.5 deg. Water tunnel results using a 3-percent scale F-18 model and a variety of wind tunnel results are used for comparison and interpretation of the flight results. The LEX vortex core breakdown point moved forward with increasing angle of attack. For a constant angle of attack, the windward LEX vortex core breakdown moves forward and inboard with sideslip and the leeward vortex core breakdown moves aft and outboard. For a constant angle of attack, the windward location of interaction moved aft with increasing sideslip and the leeward interaction moved forward.
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    Subject Category: 02
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    Report Date: January 1990
    No. Pages: 27
    Keywords:      Angle of attack; F-18 aircraft; Flow distribution; In-flight monitoring; NASA programs; Research aircraft
    Notes: AIAA Paper 90-0231. Presented at the AIAA, 28th Aerospace Sciences Meeting, Reno, Nevada, January 8-11, 1990.


  20. AN AUTOMATED CALIBRATION LABORATORY FOR FLIGHTRESEARCH INSTRUMENTATION: REQUIREMENTS AND APROPOSED DESIGN APPROACH , Technical Memorandum
    Authors: Nora Oneill-Rood and Richard D. Glover
    Report Number: NASA-TM-101719
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: NASA's Dryden Flight Research Facility (Ames-Dryden), operates a diverse fleet of research aircraft which are heavily instrumented to provide both real time data for in-flight monitoring and recorded data for postflight analysis. Ames-Dryden's existing automated calibration (AUTOCAL) laboratory is a computerized facility which tests aircraft sensors to certify accuracy for anticipated harsh flight environments. Recently, a major AUTOCAL lab upgrade was initiated; the goal of this modernization is to enhance productivity and improve configuration management for both software and test data. The new system will have multiple testing stations employing distributed processing linked by a local area network to a centralized database. The baseline requirements for the new AUTOCAL lab and the design approach being taken for its mechanization are described.
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    Subject Category: 62
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    Report Date: May 1990
    No. Pages: 16
    Keywords:      Calibrating; Data bases; Distributed processing; In-flight monitoring; Laboratories; Local area networks; Postflight analysis; Research aircraft
    Notes: Presented at the 36th ISA International Instrumentation Symposium, Denver, Colorado, May 7-10, 1990.


  21. OVERVIEW OF THE NASA AMES-DRYDEN INTEGRATED TEST FACILITY , Conference Paper
    Authors: Dale Mackall, David Mcbride and Dorothea Cohen
    Report Number: NASA-TM-101720
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An overview of the Integrated Test Facility (ITF) and the real-time systems being developed to operate it are outlined. The generic capabilities of the ITF real-time systems, the real-time data recording, and the remotely augmented vehicle (RAV) monitoring system are discussed. The benefits of applying simulation to aircraft-in-the-loop testing and the RAV monitoring system capabilities to the X-29A flight research program are considered.
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    Report Date: May 1990
    No. Pages: 15
    Keywords:      Flight simulation; Flight simulators; Ground tests; Real time operation; Research aircraft; Test facilities
    Notes: Proceedings, International Instrumentation Symposium, 36th, Denver, Colorado, May 6-10, 1990. Research Triangle Park, North Carolina, Instrument Society of America, 1990, p. 667-681


  22. EFFECTS OF SIMPLIFYING ASSUMPTIONS ON OPTIMAL TRAJECTORY ESTIMATION FOR A HIGH-PERFORMANCE AIRCRAFT , Technical Memorandum
    Authors: Lura E. Kern (NASA Ames Research Center), Steve D. Belle (PRC Systems Services Co.) and Eugene L. Duke (NASA Hugh L. Dryden Flight Research Facility)
    Report Number: NASA-TM-101721
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: When analyzing the performance of an aircraft, certain simplifying assumptions, which decrease the complexity of the problem, can often be made. The degree of accuracy required in the solution may determine the extent to which these simplifying assumptions are incorporated. A complex model may yield more accurate results if it describes the real situation more thoroughly. However, a complex model usually involves more computation time, makes the analysis more difficult, and often requires more information to do the analysis. Therefore, to choose the simplifying assumptions intelligently, it is important to know what effects the assumptions may have on the calculated performance of a vehicle. Several simplifying assumptions are examined, the effects of simplified models to those of the more complex ones are compared, and conclusions are drawn about the impact of these assumptions on flight envelope generation and optimal trajectory calculation. Models which affect an aircraft are analyzed, but the implications of simplifying the model of the aircraft itself are not studied. The examples are atmospheric models, gravitational models, different models for equations of motion, and constraint conditions.
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    Report Date: April 1990
    No. Pages: 27
    Keywords:      Aircraft performance; Atmospheric models; Computation; Estimating; Simplification; Trajectories; Trajectory optimization
    Notes: Presented at the AIAA Atmospheric Flight Mechanics Conference, Boston, MA, August 14-16, 1989


  23. FLUTTER CLEARANCE OF THE F-14A VARIABLE-SWEEPTRANSITION FLIGHT EXPERIMENT AIRPLANE, PHASE 2 , Technical Memorandum
    Authors: Lawrence C. Freudinger and Michael W. Kehoe
    Report Number: NASA-TM-101717
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An F-14A aircraft was modified for use as the test-bed aircraft for the variable-sweep transition flight experiment (VSTFE) program. The VSTFE program was a laminar flow research program designed to measure the effects of wing sweep on laminar flow. The airplane was modified by adding an upper surface foam and fiberglass glove to the right wing. An existing left wing glove had been added for the previous phase of the program. Ground vibration and flight flutter testing were accomplished to verify the absence of aeroelastic instabilities within a flight envelope of Mach 0.9 or 450 knots, calibrated airspeed, whichever was less. Flight test data indicated satisfactory damping levels and trends for the elastic structural modes of the airplane. Ground vibration test data are presented along with in-flight frequency and damping estimates, time histories and power spectral densities of in-flight sensors, and pressure distribution data.
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    Report Date: July 1990
    No. Pages: 43
    Keywords:      F-14 aircraft; Flight tests; Flutter; Laminar flow; Variable sweep wings; Aeroelasticity; Calibrating; Clearances; Damping


  24. REAL-TIME AERODYNAMIC HEATING AND SURFACE TEMPERATURE CALCULATIONS FOR HYPERSONIC FLIGHT SIMULATION , Technical Memorandum
    Authors: Robert D. Quinn and Leslie Gong
    Report Number: NASA-TM-4222
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A real-time heating algorithm was derived and installed on the Ames Research Center Dryden Flight Research Facility real-time flight simulator. This program can calculate two- and three-dimensional stagnation point surface heating rates and surface temperatures. The two-dimensional calculations can be made with or without leading-edge sweep. In addition, upper and lower surface heating rates and surface temperatures for flat plates, wedges, and cones can be calculated. Laminar or turbulent heating can be calculated, with boundary-layer transition made a function of free-stream Reynolds number and free-stream Mach number. Real-time heating rates and surface temperatures calculated for a generic hypersonic vehicle are presented and compared with more exact values computed by a batch aeroheating program. As these comparisons show, the heating algorithm used on the flight simulator calculates surface heating rates and temperatures well within the accuracy required to evaluate flight profiles for acceptable heating trajectories.
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    Subject Category: 34
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    Report Date: August 1990
    No. Pages: 44
    Keywords:      Aerodynamic heating; Algorithms; Flight simulation; Flight simulators; Hypersonic Flight; Hypersonic vehicles; Real time operation; Stagnation point; Surface temperature


  25. PROPULSION SYSTEM-FLIGHT CONTROL INTEGRATION ANDOPTIMIZATION: FLIGHT EVALUATION AND TECHNOLOGY TRANSITION , Technical Memorandum
    Authors: Frank W. Burcham, Jr., Glenn B. Gilyard and Lawrence P. Myers
    Report Number: NASA-TM-4207
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Integration of propulsion and flight control systems and their optimization offers significant performance improvements. Research programs were conducted which have developed new propulsion and flight control integration concepts, implemented designs on high-performance airplanes, demonstrated these designs in flight, and measured the performance improvements. These programs, first on the YF-12 airplane, and later on the F-15, demonstrated increased thrust, reduced fuel consumption, increased engine life, and improved airplane performance; with improvements in the 5 to 10 percent range achieved with integration and with no changes to hardware. The design, software and hardware developments, and testing requirements were shown to be practical.
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    Report Date: July 1990
    No. Pages: 26
    Keywords:      Control systems design; Electronic control; Engine control; Flight control; Propulsion system performance; Systems integration
    Notes: Presented at the AIAA 26th Joint Propulsion Conference, Orlando, Florida, July 16-18, 1990.


  26. A PRELIMINARY EVALUATION OF AN F100 ENGINE PARAMETER ESTIMATION PROCESS USING FLIGHT DATA , Technical Memorandum
    Authors: Trindel A. Maine, Glenn B. Gilyard and Heather H. Lambert
    Report Number: NASA-TM-4216
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The parameter estimation algorithm developed for the F100 engine is described. The algorithm is a two-step process. The first step consists of a Kalman filter estimation of five deterioration parameters, which model the off-nominal behavior of the engine during flight. The second step is based on a simplified steady-state model of the compact engine model (CEM). In this step, the control vector in the CEM is augmented by the deterioration parameters estimated in the first step. The results of an evaluation made using flight data from the F-15 aircraft are presented, indicating that the algorithm can provide reasonable estimates of engine variables for an advanced propulsion control law development.
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    Subject Category: 34
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    Report Date: August 1990
    No. Pages: 32
    Keywords:      Control systems design; Digital systems; Engine control; F-15 aircraft; Kalman filters; Propulsion system performance; Turbofan engines
    Notes: Presented at the 26th AIAA Joint Propulsion Conference, Orlando, Florida, July 16-18, 1990


  27. PREDICTED AND MEASURED IN-FLIGHT WING DEFORMATIONSOF A FORWARD-SWEPT-WING AIRCRAFT , Conference Paper
    Authors: William A. Lokos
    Report Number: H-1612
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An electrooptical flight-deflection measurement system (FDMS) is described in terms of its use in structural testing of the composite forward-swept wing of the X-29 aircraft. The wing deflected shapes measured using the present system are compared to the shapes predicted by NASTRAN and other codes as well as data from ground-test load measurements. The electrooptical FDMS is based on a control unit, two receivers, a target driver, and 12-16 IR LED targets. The FDMS determines the in-flight deflected wing shapes at a variety of altitudes at Mach 0.9, and the results are compared to the analytically predicted wing-twist distributions. The FDMS data describe the predicted increasing streamwise twist with increasing dynamic pressure and suggest that the streamwise twist is more prevalent at the inboard measurement station than at the wing tip. This hook shape is not represented in the predicted data, and suggestions are given for improving the modeling of the X-29 wing.
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    Report Date: January 1990
    No. Pages: 20
    Keywords:      Aeroelestic reseach wing; In-flight monitoring; Swept forward wings; Vibration measurements; Wing oscillations; X-29 aircraft
    Notes: Proceedings, Society of Flight Test Engineers, Annual Symposium, 21st, Garden Grove, California, Aug. 6-10, 1990, (A92-35926 14-01). Lancaster, California, Society of Flight Test Engineers, 1990, p. 3.1-1 to 3.1-20.


  28. FLIGHT TEST OF A TRAJECTORY CONTROLLER USING LINEARIZINGTRANSFORMATIONS WITH MEASUREMENT FEEDBACK , Conference Paper
    Authors: Robert F. Antoniewicz (NASA Hugh L. Dryden Flight Research Facility), Eugene L. Duke (NASA Flight Research Center) and P. K. A. Menon (Georgia Institute of Technology)
    Report Number: H-1619
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The design of nonlinear controllers has relied on the use of detailed aerodynamic and engine models that must be associated with the control law in the flight system implementation. Many of these controllers have been applied to vehicle flightpath control problems and have attempted to combine both inner- and outer-loop control functions in a single controller. This paper presents an alternate approach to the design of outer-loop controllers. The approach simplifies the outer-loop design problem by separating the inner-loop (stabilization and control) from the outer-loop (guidance and navigation) functions. Linearizing transformations are applied using measurement feedback to eliminate the need for detailed aircraft models in outer-loop control applications. Also discussed is an implementation of the controller. This implementation was tested on a six-degree-of-freedom F-15 simulation and in flight on an F-15 aircraft. Proof of the concept is provided by flight test data which is presented and discussed.
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    Report Date: January 1990
    No. Pages: 15
    Keywords:      Aerodynamic characteristics; F-15 aircraft; Feedback control; Flight tests; Trajectory control
    Notes: AIAA Guidance, Navigation and Control Conference, Portland, Oregon, Aug. 20-22, 1990, Technical Papers. Part 1 (A90-47576 21-08). Washington, DC, American Institute of Aeronautics and Astronautics, 1990, p. 518-532.


  29. ESTIMATING SHORT-PERIOD DYNAMICS USING ANEXTENDED KALMAN FILTER , Technical Memorandum
    Authors: Jeffrey E. Bauer (NASA Ames Research Center) and Dominick Andrisani (Purdue University)
    Report Number: NASA-TM-101722
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: An extended Kalman filter (EKF) is used to estimate the parameters of a low-order model from aircraft transient response data. The low-order model is a state space model derived from the short-period approximation of the longitudinal aircraft dynamics. The model corresponds to the pitch rate to stick force transfer function currently used in flying qualities analysis. Because of the model chosen, handling qualities information is also obtained. The parameters are estimated from flight data as well as from a six-degree-of-freedom, nonlinear simulation of the aircraft. These two estimates are then compared and the discrepancies noted. The low-order model is able to satisfactorily match both flight data and simulation data from a high-order computer simulation. The parameters obtained from the EKF analysis of flight data are compared to those obtained using frequency response analysis of the flight data. Time delays and damping ratios are compared and are in agreement. This technique demonstrates the potential to determine, in near real time, the extent of differences between computer models and the actual aircraft. Precise knowledge of these differences can help to determine the flying qualities of a test aircraft and lead to more efficient envelope expansion.
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    Report Date: January 1990
    No. Pages: 38
    Keywords:      Aircraft performance; Flight characteristics; Flight simulation; Kalman filters; Nonlinear systems; Parameter identification; State estimation; Transient response
    Notes: Presented at the 5th Biannual Flight Test Conference, Ontario, Canada, May 21-24, 1990.


  30. A PRELIMINARY EVALUATION OF AN F100 ENGINE PARAMETER ESTIMATION PROCESS USING FLIGHT DATA
    Authors: Trindel A. Maine, Glenn B. Gilyard and Heather H. Lambert
    Report Number: NASA-TM-4216
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The increasing use of digital engine control allows significant improvement in the performance of aircraft engines. This improvement can be achieved by the use of sophisticated control algorithms designed to recover the full performance potential of the propulsion system. The NASA Ames Research Center, Dryden Flight Research Facility; McDonnell Aircraft Company; and Pratt & Whitney are in the process of developing and flight testing a performance seeking control (PSC) system on the NASA F-15 research aircraft to optimize the near-steady-state performance of the F100 turbofan based propulsion system. The paper is a preliminary evaluation of the engine parameter estimation algorithm which is the primary adaptive element of the PSC algorithm. An evaluation has been made using flight data from the F-15 airplane. The flight data presented were obtained at Mach 0.90 and 30,000 ft and at three throttle positions, one of which was at intermediate power. Based on the theoretical formulation and the limited evaluation using flight data, it appears that this estimation algorithm can provide reasonable estimates of an extended set of engine variables needed for advanced propulsion control law development. However, it must be noted that conclusions drawn from this investigation are not strong because of a lack of independent flight measurements of many of the variables beings estimated. Additional sensors or independently derived estimates of many of the extended variables are needed to firmly establish the validity of the estimation algorithm.
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    Subject Category: 34
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    Report Date: August 1990
    No. Pages: 29
    Keywords:      Engine parameter estimation; F-15; Kalman filter; Performance seeking control; Propulsion system
    Notes: Prepared as a paper presented at the AIAA 26th Joint Propulsion Conference, July 16-18, 1990, Orlando, Florida.


  31. FLIGHT CONTROL SYSTEM DESIGN FACTORS FOR APPLYING AUTOMATED TESTING TECHNIQUES , Technical Memorandum
    Authors: Joel R. Sitz (NASA Hugh L. Dryden Flight Research Facility) and Todd H. Vernon (Planning Research Corp.)
    Report Number: NASA-TM-4242
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Automated validation of flight-critical embedded systems is being done at ARC Dryden Flight Research Facility. The automated testing techniques are being used to perform closed-loop validation of man-rated flight control systems. The principal design features and operational experiences of the X-29 forward-swept-wing aircraft and F-18 High Alpha Research Vehicle (HARV) automated test systems are discussed. Operationally applying automated testing techniques has accentuated flight control system features that either help or hinder the application of these techniques. The paper also discusses flight control system features which foster the use of automated testing techniques.
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    Report Date: October 1990
    No. Pages: 17
    Keywords:      Angle of attack; Automatic test equipment; Control systems design; F-18 aircraft; Flight control; Research vehicles; Swept forward wings; X-29 aircraft
    Notes: Presented at the IEEE 9th Annual Digital Avionics System Conference, Virginia Beach, Virginia, October 15-18, 1990.


  32. VALIDATION OF THE F-18 HIGH ALPHA RESEARCH VEHICLE FLIGHT CONTROL AND AVIONICS SYSTEMS MODIFICATIONS , Technical Memorandum
    Authors: Vince Chacon, Joseph W. Pahle and Victoria A. Regenie
    Report Number: NASA-TM-101723
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The verification and validation process is a critical portion of the development of a flight system. Verification, the steps taken to assure the system meets the design specification, has become a reasonably understood and straightforward process. Validation is the method used to ensure that the system design meets the needs of the project. As systems become more integrated and more critical in their functions, the validation process becomes more complex and important. The tests, tools, and techniques which are being used for the validation of the high alpha research vehicle (HARV) turning vane control system (TVCS) are discussed and the problems and their solutions are documented. The emphasis of this paper is on the validation of integrated system.
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    Report Date: October 1990
    No. Pages: 13
    Keywords:      Avionics; Control systems design; Flight control; Performance tests; Research vehicles; Vanes
    Notes: Presented at the IEEE 9th Digital Avionics Systems Conference, Virginia Beach, Virginia, October 15-18, 1990.


  33. F-18 HIGH ALPHA RESEARCH VEHICLE SURFACE PRESSURES:INITIAL IN-FLIGHT RESULTS AND CORRELATION WITH FLOW VISUALIZATIONAND WIND-TUNNEL DATA , Technical Memorandum
    Authors: David F. Fisher, Daniel W. Banks and David M. Richwine
    Report Number: NASA-TM-101724
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Pressure distributions measured on the forebody and the leading-edge extensions (LEX's) of the NASA F-18 high alpha research vehicle (HARV) were reported at 10 and 50 degree angles of attack and at Mach 0.20 to 0.60. The results were correlated with HARV flow visualization and 6-percent scale F-18 wind-tunnel-model test results. The general trend in the data from the forebody was for the maximum suction pressure peaks to first appear at an angle of attack (alpha) of approximately 19 degrees and increase in magnitude with angle of attack. The LEX pressure distribution general trend was the inward progression and increase in magnitude of the maximum suction peaks up to vortex core breakdown and then the decrease and general flattening of the pressure distribution beyond that. No significant effect of Mach number was noted for the forebody results. However, a substantial compressibility effect on the LEX's resulted in a significant reduction in vortex-induced suction pressure as Mach number increased. The forebody primary and the LEX secondary vortex separation lines, from surface flow visualization, correlated well with the end of pressure recovery, leeward and windward, respectively, of maximum suction pressure peaks. The flight to wind-tunnel correlations were generally good with some exceptions.
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    Subject Category: 02
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    Report Date: August 1990
    No. Pages: 36
    Keywords:      Angle of attack; F-18 aircraft; Flight tests; Forebodies; Leading edges; Mach number; Pressure distribution; Wind tunnel tests


  34. FLIGHT-TESTING OF THE SELF-REPAIRING FLIGHT CONTROL SYSTEM USING THE F-15 HIGHLY INTEGRATED DIGITAL ELECTRONIC CONTROL FLIGHT RESEARCH FACILITY , Technical Memorandum
    Authors: James F. Stewart and Thomas L. Shuck
    Report Number: NASA-TM-101725
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight tests conducted with the self-repairing flight control system (SRFCS) installed on the NASA F-15 highly integrated digital electronic control aircraft are described. The development leading to the current SRFCS configuration is highlighted. Key objectives of the program are outlined: (1) to flight-evaluate a control reconfiguration strategy with three types of control surface failure; (2) to evaluate a cockpit display that will inform the pilot of the maneuvering capacity of the damage aircraft; and (3) to flight-evaluate the onboard expert system maintenance diagnostics process using representative faults set to occur only under maneuvering conditions. Preliminary flight results addressing the operation of the overall system, as well as the individual technologies, are included.
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    Report Date: August 1990
    No. Pages: 17
    Keywords:      Control configured vehicles; Expert systems; F-15 aircraft; Flight control; Flight tests; Self repairing devices
    Notes: Prsented at the AIAA/SFTE/DGLR/SETP 5th Biannual Flight Test Conference, Ontario, Canada, May 21-24, 1990.


  35. A PROPOSED KALMAN FILTER ALGORITHM FOR ESTIMATION OFUNMEASURED OUTPUT VARIABLES FOR AN F100 TURBOFAN ENGINE , Technical Memorandum
    Authors: Gurbux S. Alag (PRC Systems Services Co.) and Glenn B. Gilyard (NASA Hugh L. Dryden Flight Research Facility)
    Report Number: NASA-TM-4234
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: To develop advanced control systems for optimizing aircraft engine performance, unmeasurable output variables must be estimated. The estimation has to be done in an uncertain environment and be adaptable to varying degrees of modeling errors and other variations in engine behavior over its operational life cycle. This paper represented an approach to estimate unmeasured output variables by explicitly modeling the effects of off-nominal engine behavior as biases on the measurable output variables. A state variable model accommodating off-nominal behavior is developed for the engine, and Kalman filter concepts are used to estimate the required variables. Results are presented from nonlinear engine simulation studies as well as the application of the estimation algorithm on actual flight data. The formulation presented has a wide range of application since it is not restricted or tailored to the particular application described.
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    Report Date: October 1990
    No. Pages: 32
    Keywords:      Algorithms; Control systems design; Kalman filters; State estimation; State vectors; Turbofan engines
    Notes: Presented at the AIAA/SAE/ASME/ASEE Joint Propulsion Conference, Orlando, Florida, July 16-18, 1990.


  36. REAL-TIME APPLICATION OF ADVANCED THREE-DIMENSIONAL GRAPHIC TECHNIQUES FOR RESEARCH AIRCRAFT SIMULATION , Technical Memorandum
    Authors: Steven B. Davis
    Report Number: NASA-TM-101730
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Visual aids are valuable assets to engineers for design, demonstration, and evaluation. Discussed here are a variety of advanced three-dimensional graphic techniques used to enhance the displays of test aircraft dynamics. The new software's capabilities are examined and possible future uses are considered.
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    Subject Category: 61
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    Report Date: December 1990
    No. Pages: 21
    Keywords:      Computer graphics; Computerized simulation; Display devices; Flight tests; Graphs (charts); Visual aids; X-29 aircraft


  37. A SIMPLE DYNAMIC ENGINE MODEL FOR USE IN A REAL-TIMEAIRCRAFT SIMULATION WITH THRUST VECTORING , Technical Memorandum
    Authors: Steven A. Johnson
    Report Number: NASA-TM-4240
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: A simple dynamic engine model was developed at the NASA Ames Research Center, Dryden Flight Research Facility, for use in thrust vectoring control law development and real-time aircraft simulation. The simple dynamic engine model of the F404-GE-400 engine (General Electric, Lynn, Massachusetts) operates within the aircraft simulator. It was developed using tabular data generated from a complete nonlinear dynamic engine model supplied by the manufacturer. Engine dynamics were simulated using a throttle rate limiter and low-pass filter. Included is a description of a method to account for axial thrust loss resulting from thrust vectoring. In addition, the development of the simple dynamic engine model and its incorporation into the F-18 high alpha research vehicle (HARV) thrust vectoring simulation. The simple dynamic engine model was evaluated at Mach 0.2, 35,000 ft altitude and at Mach 0.7, 35,000 ft altitude. The simple dynamic engine model is within 3 percent of the steady state response, and within 25 percent of the transient response of the complete nonlinear dynamic engine model.
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    Report Date: October 1990
    No. Pages: 21
    Keywords:      Aircraft models; F-18 aircraft; Real time operation; Scale models; Thrust distribution; Thrust vector control; Turbofan engines
    Notes: Presented at the AIAA/SAE/ASME/ASEE Joint Propulsion Conference, Orlando, Florida, July 16-18, 1990.


  38. IN-FLIGHT FLOW VISUALIZATION WITH PRESSURE MEASUREMENTS AT LOW SPEEDS ON THE NASA F-18 HIGH ALPHA RESEARCH VEHICLE , Technical Memorandum
    Authors: John H. Delfrate, David F. Fisher and Fanny A. Zuniga
    Report Number: NASA-TM-101726
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: In-flight results from surface and off-surface flow visualizations and from extensive pressure distributions document the vortical flow on the leading edge extensions (LEX) and forebody of the NASA F-18 high alpha research vehicle for low speeds and angles of attack up to 50 degs. Surface flow visualization data, obtained using the emitted fluid technique, were used to define separation lines and laminar separation bubbles. Off-surface flow visualization data, obtained by smoke injection, were used to document both the path of the vortex cores and the location of vortex core breakdown. The location of vortex core breakdown correlated well with the loss of suction pressure on the LEX and with the flow visualization results from ground facilities. Surface flow separation lines on the LEX and forebody corresponded well with the end of pressure recovery under the vortical flows. Correlation of the pressures with wind tunnel results show fair to good correlation.
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    Report Date: October 1990
    No. Pages: 45
    Keywords:      Boundary layer separation; Flight tests; Flow distribution; Flow visualization; Pressure measurement; Vortex breakdown; Vortices
    Notes: Presented at the AGARD Vortex Flow Aerodynamics Conference, Scheveningen, Netherlands, October 1-4, 1990.


  39. TECHNIQUES FOR HOT STRUCTURES TESTING , Technical Memorandum
    Authors: V. Michael Deangelis and Roger A. Fields
    Report Number: NASA-TM-101727
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Hot structures testing have been going on since the early 1960's beginning with the Mach 6, X-15 airplane. Early hot structures test programs at NASA-Ames-Dryden focused on operational testing required to support the X-15 flight test program, and early hot structures research projects focused on developing lab test techniques to simulate flight thermal profiles. More recent efforts involved numerous large and small hot structures test programs that served to develop test methods and measurement techniques to provide data that promoted the correlation of test data with results from analytical codes. In Nov. 1988 a workshop was sponsored that focused on the correlation of hot structures test data with analysis. Limited material is drawn from the workshop and a more formal documentation is provided of topics that focus on hot structures test techniques used at NASA-Ames-Dryden. Topics covered include the data acquisition and control of testing, the quartz lamp heater systems, current strain and temperature sensors, and hot structures test techniques used to simulate the flight thermal environment in the lab.
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    Report Date: November 1990
    No. Pages: 28
    Keywords:      Aerothermodynamics; Aircraft structures; Flight tests; High temperature tests; Space environment simulation; Thermal environments; Thermal stresses
    Notes: Presented at the First Thermal Structures Conference, Charlottesville, Virginia, November 13-15, 1990.


  40. INTEGRATED FLIGHT-PROPULSION CONTROL CONCEPTS FOR SUPERSONIC TRANSPORT AIRPLANES , Technical Memorandum
    Authors: Frank W. Burcham, Jr., Glenn B. Gilyard and Paul A. Gelhausen
    Report Number: NASA-TM-101728
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Integration of propulsion and flight control systems will provide significant performance improvements for supersonic transport airplanes. Increased engine thrust and reduced fuel consumption can be obtained by controlling engine stall margin as a function of flight and engine operating conditions. Improved inlet pressure recovery and decreased inlet drag can result from inlet control system integration. Using propulsion system forces and moments to augment the flight control system and airplane stability can reduce the flight control surface and tail size, weight, and drag. Special control modes may also be desirable for minimizing community noise and for emergency procedures. The overall impact of integrated controls on the takeoff gross weight for a generic high speed civil transport is presented.
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    Report Date: November 1990
    No. Pages: 18
    Keywords:      Aircraft design; Control theory; Flight control; Propulsion system performance; Supersonic transports; Systems integration
    Notes: Presented at the Society of Automotive Engineers (SAE), Inc. Aerotech Conference, Long Beach, CA, October 1-4, 1990. Prepared in cooperation with NASA-Ames Research Center, Moffett Field, CA.


  41. A PRELIMINARY LOOK AT TECHNIQUES USED TO OBTAINAIRDATA FROM FLIGHT AT HIGH ANGLES OF ATTACK , Technical Memorandum
    Authors: Timothy R. Moes and Stephen A. Whitmore
    Report Number: NASA-TM-101729
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: Flight research at high angles of attack has posed new problems for airdata measurements. New sensors and techniques for measuring the standard airdata quantities of static pressure, dynamic pressure, angle of attack, and angle of sideslip were subsequently developed. The ongoing airdata research supporting NASA's F-18 high alpha research program is updated. Included are the techniques used and the preliminary results. The F-18 aircraft was flown with three research airdata systems: a standard airdata probe on the right wingtip, a self-aligning airdata probe on the left wingtip, and a flush airdata system on the nose cone. The primary research goal was to obtain steady-state calibrations for each airdata system up to an angle of attack of 50 deg. This goal was accomplished and preliminary accuracies of the three airdata systems were assessed and are presented. An effort to improve the fidelity of the airdata measurements during dynamic maneuvering is also discussed. This involved enhancement of the aerodynamic data with data obtained from linear accelerometers, rate gyros, and attitude gyros. Preliminary results of this technique are presented.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 06
    Availability:
        Format(s) on-line:
          PDF (2,178 KBytes)
    Report Date: December 1990
    No. Pages: 51
    Keywords:      Accelerometers; Aerodynamics; Angle of attack; Attitude gyros; Calibrating; Dynamic pressure; F-18 aircraft; Sideslip; Static pressure
    Notes: Presented at the High Angle of Attack Technology Symposium, Hampton, Virginia, October 30 - November 1, 1990.


  42. MONITORING TECHNIQUES FOR THE X-29A AIRCRAFT'SHIGH-SPEED ROTATING POWER TAKEOFF SHAFT , Technical Memorandum
    Authors: David F. Voracek
    Report Number: NASA-TM-101731
    Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
    Abstract: The experimental X-29A forward swept-wing aircraft has many unique and critical systems that require constant monitoring during ground or flight operation. One such system is the power takeoff shaft, which is the mechanical link between the engine and the aircraft-mounted accessory drive. The X-29A power takeoff shaft opertes in a range between 0 and 16,810 rpm, is longer than most jet engine power takeoff shafts, and is made of graphite epoxy material. Since the X-29A aircraft operates on a single engine, failure of the shaft during flight could lead to loss of the aircraft. The monitoring techniques and test methods used during power takeoff shaft ground and flight operations are discussed. Test data are presented in two case studies where monitoring and testing of the shaft dynamics proved instrumental in discovering and isolating X-29A power takeoff shaft problems. The first study concerns the installation of an unbalanced shaft. The effect of the unbalance on the shaft vibration data and the procedure used to correct the problem are discussed. The second study deals with the shaft exceeding the established vibration limits during flight. This case study found that the vibration of connected rotating machinery unbalances contributed to the excessive vibration level of the shaft. The procedures used to identify the contributions of other rotating machinery unbalances to the power takeoff shaft unbalance are discussed.
    Distribution/Availability: Unclassified - Unlimited
    Subject Category: 05
    Availability:
        Format(s) on-line:
          PDF (925 KBytes)
    Report Date: December 1990
    No. Pages: 25
    Keywords:      Engine failure; Flight operations; Graphite-epoxy composites; Ground operational support system; High speed; Rotating shafts; Takeoff; Vibration; X-29 aircraft
    Notes: Presented at the 2nd International Machinery Monitoring and Diagnostic Conference, Los Angeles, CA, October 22-25, 1990.
 
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