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- PILOT-INDUCED OSCILLATION RESEARCH: STATUS AT THE END OF THE CENTURY , Conference Publication
Authors: Mary F. Shafer and Paul Steinmetz
Report Number: NASA-CP-2001-210389VOL1-3
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The workshop "Pilot-Induced Oscillation Research: The Status at the End of the Century," was held at NASA Dryden Flight Research Center on 6 to 8 April 1999. The presentations at this conference addressed the most current information available, addressing regulatory issues, flight test, safety, modeling, prediction, simulation, mitigation or prevention, and areas that require further research. All presentations were approved for publication as unclassified documents with no limits on their distribution. This proceedings includes the viewgraphs (some with author's notes) used for thirty presentations that were actually given and two presentations that were not given because of time limitations. Four technical papers on this subject are also included.
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Report Date: April 2001
No. Pages: 477
Funding Organization: 529-55-24-E8-RR-00-000
Keywords: Flight control; Flight safety; Pilot-induced oscillation; Simulation of flight test
Notes: Presentations from workshop held at NASA Dryden Flight Research Center, 6 to 8 April, 1999. Errata, May 1, 2001. Pages 413-438, which were incomplete at time of printing, are now available and have been replaced on the pdf.
- FLIGHT DEMONSTRATION OF X-33 VEHICLE HEALTH MANAGEMENT SYSTEM COMPONENTS ON THE F/A-18 SYSTEMS RESEARCH AIRCRAFT , Technical Memorandum
Authors: Keith A. Schweikhard, W. Lance Richards, John Theisen, William Mouyos and Raymond Garbos
Report Number: NASA-TM-2001-209037
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The X-33 reusable launch vehicle demonstrator has identified the need to implement a vehicle health monitoring system that can acquire data that monitors system health and performance. Sanders, a Lockheed Martin Company, has designed and developed a COTS-based open architecture system that implements a number of technologies that have not been previously used in a flight environment. NASA Dryden Flight Research Center and Sanders teamed to demonstrate that the distributed remote health nodes, fiber optic distributed strain sensor, and fiber distributed data interface communications components of the X- 33 vehicle health management (VHM) system could be successfully integrated and flown on a NASA F-18 aircraft. This paper briefly describes components of X-33 VHM architecture flown at Dryden and summarizes the integration and flight demonstration of these X-33 VHM components. Finally, it presents early results from the integration and flight efforts.
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Subject Category: 06
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Report Date: December 2001
No. Pages: 15
Funding Organization: 529-35-34-E8-RR-00-000
Keywords: Aircraft systems; Bragg gratings; Fiber optic sensing; Flight testing; Integrated vehicle health management; Intelligent systems; IVHM; Structural health monitoring
Notes: Presented at the 19th Digital Avionics Systems Conference, Philadelphia, Pennsylvania on October 7-11, 2000.
- DEVELOPMENT OF A MARS AIRPLANE ENTRY, DESCENT, AND FLIGHT TRAJECTORY , Technical Memorandum
Authors: James E. Murray and Paul V. Tartabini
Report Number: NASA-TM-2001-209035
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: An entry, descent, and flight (EDF) trajectory profile for a Mars airplane mission is defined as consisting of the following elements: ballistic entry of an aeroshell; supersonic deployment of a decelerator parachute; subsonic release of a heatshield; release, unfolding, and orientation of an airplane to flight attitude; and execution of a pullup maneuver to achieve trimmed, horizontal flight. Using the Program to Optimize Simulated Trajectories (POST) a trajectory optimization problem was formulated. Model data representative of a specific Mars airplane configuration, current models of the Mars surface topography and atmosphere, and current estimates of the interplanetary trajectory, were incorporated into the analysis. The goal is to develop an EDF trajectory to maximize the surface-relative altitude of the airplane at the end of a pullup maneuver, while subject to the mission design constraints. The trajectory performance was evaluated for three
potential mission sites and was found to be site-sensitive. The trajectory performance, examined for sensitivity to a number of design and constraint variables, was found to be most sensitive to airplane mass, aerodynamic performance characteristics, and the pullup Mach constraint. Based on the results of this sensitivity study, an airplane-drag optimized trajectory was developed that showed a significant performance improvement.
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Report Date: January 2001
No. Pages: 20
Keywords: Aircraft design; Atmospheric entry; Flight mechanics; Mars probes; Trajectory optimization
Notes: Presented at 39th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 8-11, 2001, AIAA-2001-0839. Paul V. Tartabini, NASA Langley Research Center, Hampton, Virginia.
- WIND-TUNNEL INVESTIGATIONS OF BLUNT-BODY DRAG REDUCTION USING FOREBODY SURFACE ROUGHNESS , Technical Memorandum
Authors: Stephen A. Whitmore, Stephanie Sprague and Jonathan W. Naughton
Report Number: NASA-TM-2001-210390
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: This paper presents results of wind-tunnel tests that demonstrate a novel drag reduction technique for blunt-based vehicles. For these tests, the forebody roughness of a blunt-based model was modified using micomachined surface overlays. As forebody roughness increases, boundary layer at the model aft thickens and reduces the shearing effect of external flow on the separated flow behind the base region, resulting in reduced base drag. For vehicle configurations with large base drag, existing data predict that a small increment in forebody friction drag will result in a relatively large decrease in base drag. If the added increment in forebody skin drag is optimized with respect to base drag, reducing the total drag of the configuration is possible. The wind-tunnel tests results conclusively demonstrate the existence of a forebody drag-base drag optimal point. The data demonstrate that the base drag coefficient corresponding to the drag minimum lies
between 0.225 and 0.275, referenced to the base area. Most importantly, the data show a drag reduction of approximately 15 percent when the drag optimum is reached. When this drag reduction is scaled to the X-33 base area, drag savings approaching 45,000 N (10,000 lbf ) can be realized.
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Subject Category: 05
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Report Date: January 2001
No. Pages: 33
Keywords: Base drag; Drag reduction; Reusable launch vehicle; Skin friction; Wind tunnel
Notes: Presented at 39th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada, January 8-11, 2001, AIAA-2001-0252. Stephanie Sprague, University of Kansas. Jonathan W. Naughton, University of Wyoming.
- NONSTATIONARY DYNAMICS DATA ANALYSIS WITH WAVELET-SVD FILTERING , Technical Memorandum
Authors: Marty Brenner and Dale Groutage
Report Number: NASA-TM-2001-210391
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Nonstationary time-frequency analysis is used for identification and classification of aeroelastic and aeroservoelastic dynamics. Time-frequency multiscale wavelet processing generates discrete energy density distributions. The distributions are processed using the singular value decomposition (SVD). Discrete density functions derived from the SVD generate moments that detect the principal features in the data. The SVD standard basis vectors are applied and then compared with a transformed-SVD, or TSVD, which reduces the number of features into more compact energy density concentrations. Finally, from the feature extraction, wavelet-based modal parameter estimation is applied.
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Subject Category: 66
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Report Date: April 2001
No. Pages: 22
Funding Organization: 529-50-04-00-RR-00-000
Keywords: Aeroelasticity; Aeroservoelasticity; Singular value decomposition; System identification; Time-frequency analysis; Wavelet analysis
Notes: Modified from paper presented at 42nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, Seattle, Washington, April 16-19, 2001. Dale Groutage, Naval Surface Warfare Center (NSWC), Bremerton, Washington.
- FOREBODY AERODYNAMICS OF THE F-18 HIGH ALPHA RESEARCH VEHICLE WITH ACTUATED FOREBODY STRAKES , Conference Paper
Authors: David F. Fisher and Daniel G. Muri
Report Number: NASA-MP-69-P-45
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Extensive pressure measurements and off-surface flow visualization were obtained on the forebody and strakes of the NASA F-18 High Alpha Research Vehicle (HARV) equipped with actuated forebody strakes.Forebody yawing moments were obtained by integrating the circumferential pressures on the forebody and strakes. Results show that large yawing moments can be generated with forebody strakes. At a 50 deg-angle-of-attack, deflecting one strake at a time resulted in a forebody yawing moment control reversal for small strake deflection angles. However, deflecting the strakes differentially about a 20 deg symmetric strake deployment eliminated the control reversal and produced a near linear variation of forebody yawing moment with differential strake deflection. At an angle of attack of 50 deg and for 0 deg and 20 deg symmetric strake deployments, a larger forebody yawing moment was generated by the forward fuselage (between the radome and the apex of the
leading-edge extensions) than on the radome where the actuated forebody strakes were located. Cutouts on the flight vehicle strakes that were not on the wind tunnel models are believed to be responsible for deficits in the suction peaks on the flight radome pressure distributions and differences in the forebody yawing moments.
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Subject Category: 02
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Report Date: May 2001
No. Pages: 16
Funding Organization: 710-55-04-E8-RR-00-000
Keywords: Aircraft; F-18; Flight; Flight Test; Flow visualization; Forebody; Smoke; Strake;Vortex; Vortical flow
Notes: Presented at the Symposium on Advanced Flow Management, Loen, Norway, May 7-10, 2001.
- DEVELOPING UNCERTAINTY MODELS FOR ROBUST FLUTTER ANALYSIS USING GROUND VIBRATION TEST DATA , Technical Memorandum
Authors: Starr Potter and Rick Lind
Report Number: NASA-TM-2001-210392
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A ground vibration test can be used to obtain information about structural dynamics that is important for flutter analysis. Traditionally, this information - such as natural frequencies of modes - is used to update analytical models used to predict flutter speeds. The ground vibration test can also be used to obtain uncertainty models, such as natural frequencies and their associated variations, that can update analytical models for the purpose of predicting robust flutter speeds. Analyzing test data using the infinity-norm, rather than the traditional 2-norm, is shown to lead to a minimum-size uncertainty description and, consequently, a least-conservative robust flutter speed. This approach is demonstrated using ground vibration test data for the Aerostructures Test Wing. Different norms are used to formulate uncertainty models and their associated robust flutter speeds to evaluate which norm is least conservative.
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Subject Category: 05
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Report Date: April 2001
No. Pages: 21
Funding Organization: 274-00-00-E8-RR-00-DDF
Keywords: Flutter; Ground vibration test; Uncertainty
Notes: Presented at 42nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference and Exhibit, Seattle, Washington, April 16-19, 2001, AIAA-2001-1585.
- THERMOSTRUCTURAL ANALYSIS OF UNCONVENTIONAL WING STRUCTURES OF A HYPER-X HYPERSONIC FLIGHT RESEARCH VEHICLE FOR THE MACH 7 MISSION , Technical Publication
Authors: William L. Ko and Leslie Gong
Report Number: NASA-TP-210398
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Heat transfer, thermal stresses, and thermal buckling analyses were performed on the unconventional wing structures of a Hyper-X hypersonic flight research vehicle (designated as X-43) subjected to nominal Mach 7 aerodynamic heating. A wing midspan cross section was selected for the heat transfer and thermal stress analyses. Thermal buckling analysis was performed on three regions of the wing skin (lower or upper); 1) a fore wing panel, 2) an aft wing panel, and 3) a unit panel at the middle of the aft wing panel. A fourth thermal buckling analysis was performed on a midspan wing segment. The unit panel region is identified as the potential thermal buckling initiation zone. Therefore, thermal buckling analysis of the Hyper-X wing panels could be reduced to the thermal buckling analysis of that unit panel. "Buckling temperature magnification factors" were established. Structural temperature-time histories are presented. The results show that the
concerns of shear failure at wing and spar welded sites, and of thermal buckling of Hyper-X wing panels, may not arise under Mach 7 conditions.
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Subject Category: 39
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Report Date: October 2001
No. Pages: 42
Funding Organization: 710-55-24-E8-RR-00-000
Keywords: Buckling temperature magnification factors; Hyper-X wing; Irregular panels; Non-uniform temperature loading; Thermal buckling
- COMPARISON OF X-31 FLIGHT AND GROUND-BASED YAWING MOMENT ASYMMETRIES AT HIGH ANGLES OF ATTACK , Technical Memorandum
Authors: Brent R. Cobleigh and Mark A. Croom
Report Number: NASA-TM-2001-210393
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 caused position saturations of the thrust-vectoring vanes and trailing-edge flaps during some stability-axis rolling maneuvers at high angles of attack. The two test aircraft had different asymmetry characteristics, and ship 2 has asymmetries that vary as a function of Reynolds number. Several aerodynamic modifications have been made to the X-31 forebody with the goal of minimizing the asymmetry. These modifications include adding transition strips on the forebody and noseboom, using two different length strakes, and increasing nose bluntness. Ultimately, a combination of forebody strakes, nose blunting, and noseboom transition strips reduced the yawing moment asymmetry enough to fully expand the high-angle-of- attack envelope. Analysis of the X-31 flight data is reviewed and compared to wind- tunnel
and water-tunnel measurements. Several lessons learned are outlined regarding high-angle-of-attack configuration design and ground testing.
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Subject Category: 05
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Report Date: May 2001
No. Pages: 18
Funding Organization: 710-55-24-E8-RR-00-000
Keywords: Asymmetry; Water tunnel test; Wind tunnel test; X-31 aircraft; Yawing moment
Notes: Paper MP-69-P-42, presented at the Symposium on Advanced Flow Management, May 7-11, 2001, Loen, Norway. Mark Croom, NASA LaRC, Hampton, VA.
- THE F-15B PROPULSION FLIGHT TEST FIXTURE: A NEW FLIGHT FACILITY FOR PROPULSION RESEARCH , Technical Memorandum
Authors: Stephen Corda, M. Jake Vachon, Nathan Palumbo, Corey Diebler, Ting Tseng, Anthony Ginn and David Richwine
Report Number: NASA-TM-2001-210395
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: The design and development of the F-15B Propulsion Flight Test Fixture (PFTF), a new facility for propulsion flight research, is described. Mounted underneath an F- 15B fuselage, the PFTF provides volume for experiment systems and attachment points for propulsion devices. A unique feature of the PFTF is the incorporation of a six-degree-of-freedom force balance. Three-axis forces and moments can be measured in flight for experiments mounted to the force balance. The NASA F-15B airplane is described, including its performance and capabilities as a research test bed aircraft. The detailed description of the PFTF includes the geometry, internal layout and volume, force-balance operation, available instrumentation, and allowable experiment size and weight. The aerodynamic, stability and control, and structural designs of the PFTF are discussed, including results from aerodynamic computational fluid dynamic calculations and structural analyses. Details
of current and future propulsion flight experiments are discussed. Information about the integration of propulsion flight experiments is provided for the potential PFTF user.
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Subject Category: 07
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Report Date: July 2001
No. Pages: 18
Funding Organization: 529-35-14-00-38-00-F-15
Keywords: F-15B flight testing; In-flight force balance; Propulsion flight test fixture; Propulsion flight testing; Rocket-based combined cycle propulsion
Notes: Presented at the 37th AIAA/SAE/ASME/ASEE Joint Propulsion Conference & Exhibit, July 8-11, 2001, Salt Lake City, Utah, AIAA 2001-3303.
- RECONFIGURABLE CONTROL DESIGN FOR THE FULL X-33 FLIGHT ENVELOPE , Technical Memorandum
Authors: M. Christopher Cotting and John J. Burken
Report Number: NASA-TM-2001-210396
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A reconfigurable control law for the full X-33 flight envelope has been designed to accommodate a failed control surface and redistribute the control effort among the remaining working surfaces to retain satisfactory stability and performance. An offline nonlinear constrained optimization approach has been used for the X-33 reconfigurable control design method. Using a nonlinear, six-degree-of-freedom simulation, three example failures are evaluated: ascent with a left body flap jammed at maximum deflection; entry with a right inboard elevon jammed at maximum deflection; and landing with a left rudder jammed at maximum deflection. Failure detection and identification are accomplished in the actuator controller. Failure response comparisons between the nominal control mixer and the reconfigurable control subsystem (mixer) show the benefits of reconfiguration. Single aerosurface jamming failures are considered. The cases evaluated are representative
of the study conducted to prove the adequate and safe performance of the reconfigurable control mixer throughout the full flight envelope. The X-33 flight control system incorporates reconfigurable flight control in the existing baseline system.
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Subject Category: 08
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Report Date: August 2001
No. Pages: 11
Funding Organization: 715-33-02-E8-23-00-T19
Keywords: Control System Reconfiguration; Emergency Backup System; Reusable Launch
Notes: Presented at AIAA Guidance, Navigation & Control Conference, Montreal, Quebec, Canada, August 6-10, 2001, AIAA-2001-4379.
- A MONTE CARLO DISPERSION ANALYSIS OF THE X-33 SIMULATION SOFTWARE , Conference paper
Authors: Peggy S. Williams
Report Number: H-2460
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: A Monte Carlo dispersion analysis has been completed on the X-33 software simulation. The simulation is based on a preliminary version of the software and is primarily used in an effort to define and refine how a Monte Carlo dispersion analysis would have been done on the final flight-ready version of the software. This report gives an overview of the processes used in the implementation of the dispersions and describes the methods used to accomplish the Monte Carlo analysis. Selected results from 1000 Monte Carlo runs are presented with suggestions for improvements in future work.
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Report Date: August 2001
No. Pages: 10
Funding Organization: 715-33-02-E8-23-00-NTA
Keywords: n/a
- AEROSERVOELASTIC UNCERTAINTY MODEL IDENTIFICATION FROM FLIGHT DATA , Technical Memorandum
Authors: Martin J. Brenner
Report Number: NASA-TM-210397
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Uncertainty modeling is a critical element in the estimation of robust stability margins for stability boundary prediction and robust flight control system development. There has been a serious deficiency to date in aeroservoelastic data analysis with attention to uncertainty modeling. Uncertainty can be estimated from flight data using both parametric and nonparametric identification techniques. The model validation problem addressed in this paper is to identify aeroservoelastic models with associated uncertainty structures from a limited amount of controlled excitation inputs over an extensive flight envelope. The challenge to this problem is to update analytical models from flight data estimates while also deriving non-conservative uncertainty descriptions consistent with the flight data. Multisine control surface command inputs and control system feedbacks are used as signals in a wavelet-based modal parameter estimation procedure for model
updates. Transfer function estimates are incorporated in a robust minimax estimation scheme to get input-output parameters and error bounds consistent with the data and model structure. Uncertainty estimates derived from the data in this manner provide an appropriate and relevant representation for model development and robust stability analysis. This model-plus-uncertainty identification procedure is applied to aeroservoelastic flight data from the NASA Dryden Flight Research Center F-18 Systems Research Aircraft.
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Subject Category: 08
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Report Date: July 2001
No. Pages: 18
Funding Organization: 529-61-14-E8-14-00-AAW
Keywords: Aeroservoelasticity; Modal identification; Morlet wavelet; Robust minimax
Notes: Presented at CEAS/AIAA International Forum on Aeroelasticity and Structural Dynamics, Madrid, Spain, June 5-7, 2001. Self-published PDF provided by author.
- DATA SYNCHRONIZATION DISCREPANCIES IN A FORMATION FLIGHT CONTROL SYSTEM , Technical Memorandum
Authors: Jack Ryan, Curtis E. Hanson, Ken A. Norlin and Michael Allen
Report Number: NASA-TM-2001-210720
Performing Organization: NASA Dryden Flight Research Center, Edwards, CA
Abstract: Aircraft hardware-in-the-loop simulation is an invaluable tool to flight test engineers; it reveals design and implementation flaws while operating in a controlled environment. Engineers, however, must always be skeptical of the results and analyze them within their proper context. Engineers must carefully ascertain whether an anomaly that occurs in the simulation will also occur in flight. This report presents a chronology illustrating how misleading simulation timing problems led to the implementation of an overly complex position data synchronization guidance algorithm in place of a simpler one. The report illustrates problems caused by the complex algorithm and how the simpler algorithm was chosen in the end. Brief descriptions of the project objectives, approach, and simulation are presented. The misleading simulation results and the conclusions then drawn are presented. The complex and simple guidance algorithms are presented with flight data
illustrating their relative success.
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Subject Category: 04
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Report Date: November 2001
No. Pages: 15
Funding Organization: 710-70-14-E8-28-00-AFF
Keywords: Global Positioning System; Inertial navigation; Automatic control; Flight simulation; Universal time
Notes: Presented at the 32nd Annual Society of Flight Test Engineers Symposium, 10-14 September 2001.
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