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Parameter-Free Data Reduction for Short-Time Hypersonic Heat Flux Measurements

JOURNAL OF THERMOPHYSICS AND HEAT TRANSFER 2025年第2期39卷 347-354页

作者： Frankel, J. I. Shu, F. Gross, A. New Mexico State Univ Mech & Aerosp Engn Dept Las Cruces NM 88003 USA

High-speed flows associated with hypersonics require significant, diverse, and successful ground-test campaigns before engaging in a flight test program. The present investigation focuses on demonstrating a new data reduction pathway for estimating the source heat flux using an elegant, though fragile, thin-film temperature gauge. The rapid response time of such sensors makes them well-suited for shock tunnel experiments. In such experiments, contaminated flow, composed of diaphragm debris, can often damage the sensing element. For the moment, the existence of an ideal coating (or even a hardened sensing element) is presumed that preserves the sensor's survivability and accuracy over several test runs. A dynamic, physics-based data reduction methodology is proposed that accounts for the protective coating and that is applicable to millisecond testing times. This paper addresses three critical issues for acquiring the best estimate of the source heat flux. These issues include a) the formation of a baseline data reduction view that accounts for thin coatings;b) the development of a dynamic calibration method that removes the need for specifying thermophysical or geometrical properties;and c) the introduction of a time-scaling that allows for calibration to be performed in alternative and convenient time frames. This last point demonstrates the concept of the dimensionless time map and its novel consequences. The procedure is demonstrated on a reduced formulation, illustrating the merit of the concepts.

关键词： Data Processing Heat Flux Rate Sensors Thin Films Shock Tunnels Aircraft Operations and Technology flight test programme High Speed Flows Thermal Storage Inverse Heat Conduction

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Investigation of Energy-Maneuverability Methods for Helicopter Performance Assessments

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JOURNAL OF AIRCRAFT 2025年第1期62卷 22-33页

作者： Tritschler, John K. Juhasz, Ondrej O'Connor, John C. Mccue, J. J. Holder, John US Naval Test Pilot Sch Patuxent River Maryland NY 20670 USA US Naval Acad Annapolis MD 21402 USA

Energy-maneuverability diagrams are an important tool that operational pilots use to understand helicopter maneuver performance through excess available power across a wide range of conditions. These representations are based upon a number of assumptions that have not been rigorously investigated for helicopter applications. The present work reports the results of an investigation into the theory and application of helicopter maneuverability through simulation and flight test. The computational portion of the work focused on a systematic investigation into some of the key simplifying assumptions that are commonly applied in the creation of energy-maneuverability diagrams using two rotor wake aerodynamic modeling methods. The flight test portion of the work provided important operational context for understanding the practical application of the simulation results. The study illustrated that the fundamental assumption employed in estimating maneuver power requirements for energy-maneuverability representations appears to be reasonable in conditions of the greatest practical relevance;however, another key assumption that is invoked to convert excess power into climb performance would likely lead to overestimating the vehicle capability in important operational conditions. Additionally, the flight test data demonstrated that, at high angles of bank, energy-maneuverability results should be considered for trending information rather than for detailed climb performance values.

关键词： Blade Loading flight test programme Energy Maneuverability Theory Energy/Maneuverability rotorcraft

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Real-Time Frequency Response Estimation from flight Data

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2012年第5期35卷 1406-1417页

作者： Holzel, Matthew S. Morelli, Eugene A. Univ Michigan Ann Arbor MI 48105 USA NASA Langley Res Ctr Dynam Syst & Control Branch Hampton VA 23681 USA

A nonparametric method for estimating frequency responses in real time based on recursive least squares in the time domain was developed and studied. The method uses sinusoidal functions at selected frequencies known to be contained in the input in conjunction with recursive least squares to rapidly estimate Fourier series coefficients for input and output time series, and thereby estimate frequency responses. Practical problems that arise when applying the discrete Fourier transform in recursive form were identified using simple examples. A general expression for accurate real-time calculation of the covariance matrix for recursive least-squares parameter estimation was developed and used to calculate valid uncertainty bounds for real-time frequency response estimates. Simulation data generated with optimized multisine inputs were used to investigate the accuracy of the proposed method for real-time frequency response estimation and to validate the calculated error bounds. The approach was also applied to flight data from a subscale jet transport aircraft. Comparisons of real-time frequency response estimates and error bounds with results from conventional postflight batch analysis showed that results from the real-time method were in statistical agreement with postflight batch estimates, with valid error bounds that properly indicated the quality of the real-time frequency response estimates.

关键词： Frequency Response Functions flight Data Analysis Control Surfaces Elevator Deflection Bode Plot Computing flight test programme Probability Density Functions Transport Aircraft Pilot

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Development and flight testing of quantitative feedback theory pitch rate stability augmentation system

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 1996年第5期19卷 1109-1115页

作者： Fontenrose, PL Hall, CE North Carolina State Univ Raleigh NC USA

The design of a quantitative feedback theory control system to improve the flight handling qualities of an unmanned aerial vehicle (UAV) at relaxed stabilities is illustrated, Aircraft flown at low static margins are significantly more efficient than conventional designs, but without stability augmentation they are difficult if not impossible to control. Longitudinal stability augmentation is achieved by using an onboard computer system to calculate control surface deflection based on the sensed pitch rate of the aircraft and pilot elevator commands. The quantitative feedback theory controller was implemented on this computer system to provide acceptable handling characteristics for a large portion of the UAV's flight envelope, The flight tests of the UAV, which resulted in the successful completion of an entire flight at a 2.5% static margin, are also described.

关键词： flight test programme Stability Augmentation Quantitative Feedback Theory Elevator Deflection Single Input Single Output Control Surfaces Aircraft Pilot Cropped Delta Wing UAV Handling Qualities

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Nonlinear model development from flight-test data for F/A-18E super hornet

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JOURNAL OF AIRCRAFT 2004年第4期41卷 692-702页

作者： Paris, AC Bonner, M Sci Applicat Int Corp Lexington Pk MD 20653 USA Naval Air Syst Command Patuxent River MD 20670 USA

Applied procedures for nonlinear aerodynamic model development and extraction from flight data for the F/A-18E Super Hornet aircraft are addressed. Parameter identification maneuver design, flight-test support technique, and aerodynamic model development are included. All-axis 3-2-1-1 maneuvers were designed to best excite the longitudinal and lateral-directional system natural frequencies of interest. The work was in support of leading edge flap degraded modes studies for the F/A-18E aircraft investigated by the Naval Air Systems Command. In this flight-test program, the effect of off-schedule, differentially frozen, leading-edge flaps was examined with the aircraft in power-approach configuration. A variety of processes was employed to develop fully nonlinear longitudinal and lateral-directional aerodynamic incremental model corrections. This effort included total aerodynamic coefficient reconstruction from flight-test data and equation error analysis for model structure development/estimation. The available data allowed for the successful extraction of incremental coefficients for aerodynamic forces and moments that resulted in an engineering simulation with increased aerodynamic fidelity in the configurations examined.

关键词： flight test programme Aerodynamic Coefficients flight Data Leading Edge Flap Angle of Attack Angle of Sideslip Aircraft Configurations test Pilots Yaw Damper Transient Response

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ROLLING TAIL DESIGN AND BEHAVIOR AS AFFECTED BY ACTUATOR HINGE MOMENT LIMITS

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JOURNAL OF AIRCRAFT 1979年第8期16卷 527-531页

作者： BALL, JN Calspan Corporation Los Angeles Calif.

Experience with the B-1 airplane has illustrated some problems that can be encountered with rolling tails. Stabilizer actuators large enough for pitch control can generate large tail rolling moments and high loads in the tail support structure. At high speeds, undesirable behavior of the flight control system can be encountered if the actuator hinge moment limit is reached. The measures used to avoid or deal with such problems on the B-1 are explained. The paper is presented as a case history to show the various interrelated consequences arising from the design decision to use a rolling tail, with particular attention to the effects of the actuator hinge moment limit.

关键词： Aerodynamic Characteristics Actuators Supersonic flight flight Control System flight test programme Angle of Attack Airframes Control Surfaces Hydraulic Actuators Linkages

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OPTIMAL LANDING OF A HELICOPTER IN AUTOROTATION

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 1988年第1期11卷 7-12页

作者： LEE, AY BRYSON, AE HINDSON, WS STANFORD UNIV DEPT AERONAUT & ASTRONAUTSTANFORDCA 94305 STANFORD UNIV DEPT MECH ENGNSTANFORDCA 94305

The landing of a helicopter in autorotation is formulated as a nonlinear optimal control problem with inequality constraints. The performance index is a weighted sum of the squares of vertical and horizontal velocity at touchdown. The control inequality constraint is a limit on the rotor thrust coefficient. The state inequality constraint is a limit on vertical sink rate. Optimal trajectories are calculated for initial conditions well within the height-velocity (H-V) restriction curve, with the helicopter in hover or forward flight. The optimal control history is similar to those used by helicopter pilots in autorotational landings. The study indicates that, subject to pilot acceptability, a substantial reduction could be made in the H-V restriction zone using optimal control techniques.

关键词： Autorotation Rotor Systems Optimal Control Pilot flight test programme Blade Element Theory flight Data Collective Pitch Control Vortex Ring State Power Coefficient

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Investigating the role of rate limiting in pilot-induced oscillations

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JOURNAL OF GUIDANCE CONTROL AND DYNAMICS 2004年第5期27卷 804-813页

作者： Klyde, DH Mitchell, DG Syst Technol Inc Hawthorne CA 90250 USA Hoh Aeronaut Inc Lomita CA 90717 USA

From the Wright Flyer to fly-by-wire, the phenomenon of pilot-induced oscillations (PIO) has been observed on prototype, experimental, and operational military and commercial aircraft. The introduction of irreversible control systems with surfaces driven by powered actuators brought many benefits along with increased system complexity and the introduction of additional nonlinearities. Chief among these nonlinearities are the hardware and software rate limits associated with the control surface actuators. Basic sizing tradeoffs conducted in the design process set the maximum rate of an actuator, whereas software rate limits are introduced to prevent overdriving the control surface when loads or structural limitations exist. It is demonstrated that, when operating as intended, there are usually no ill effects associated with rate limits;however, certain conditions can lead to a highly saturated condition. This results in the sudden introduction of significant added phase lags to the pilot-vehicle system. In many cases, the end result is often PIO or other related loss of control events. One of the earliest well-documented PIO events involving rate limiting occurred on the first flight of the X-15. This event is significant in that common linear systems analysis techniques do not reveal a susceptibility to PIO. The analysis of X-15 flight 1-1-5 and the results of a more recent flight research program are presented in detail.

关键词： Pilot Induced Oscillation Nonlinear Actuators Aircraft Pilot Commercial Aircraft Describing Function Flying Wires Handling Qualities flight test programme flight Control System Airframes

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Advances in real-time aerodynamic model identification

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JOURNAL OF AIRCRAFT 2005年第1期42卷 73-79页

作者： Hui, K Auriti, L Ricciardi, J Natl Res Council Canada Flight Res Lab Inst Aerosp Res Ottawa ON K1A 0R6 Canada

This paper describes the development of an in-flight, near real-time, aerodynamic model identification (RAMI) technique, for fixed-wing aircraft. This technique can be used to develop and validate an aerodynamic model to the Federal Aviation Administration AC 120 Level D flight simulator aerodynamic model criteria;and initial demonstrations of its capabilities are taking place on the National Research Council Falcon 20 research aircraft. As part of the RAMI development, innovative, self-contained software algorithms and hardware processing have been designed for onboard use. As well, novel hardware and software techniques are being applied to calibrate fundamental in-flight measurements, such as air data, in real time. The RAMI technique uses the "2-3-1-1" (alternating control input steps in seconds) maneuver, executed in each of the pitch, roll and yaw axes. These maneuvers are flown in flight conditions that span the entire flight envelope of interest. At each flight condition, the small perturbation stability and control derivatives (point model) are determined in real time. An automatic curve-fit process then combines the point models to develop a global aerodynamic model of the aircraft. In addition, an online model validation process is executed to confirm the integrity of the developed model, by comparing the model response to flight data. Excellent aircraft model fidelity has been demonstrated using the RAMI technique.

关键词： Aerodynamic Simulation Yaw Damper flight Data Federal Aviation Administration Fixed Wing Aircraft flight Envelope flight Simulators flight test programme Six Degree of Freedom FORTRAN

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KEEP EAGLE F-15E high angle-of-attack flight test program

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JOURNAL OF AIRCRAFT 1997年第3期34卷 265-270页

作者： Sitz, DM Nelson, DC Carpenter, MR MCDONNELL DOUGLAS CORP AEROSPDEPT 281EDWARDS AFBCA 93523

The KEEP EAGLE night test program was conducted from August 1994 until August 1995 at Edwards Air Force Base by a combined government/contractor test team to evaluate improvements to F-15E high angle-of-attack and spin recovery characteristics. This paper traces the program from its inception in 1992 until its conclusion in 1995, with emphasis on the test approach and flight test techniques employed for this high-risk program, Specifically, the test approach included novel assessments of spin recovery control power early in the night test program using controlled buildups in yaw rate. The program also used simulation effectively to improve test efficiency and maintain test team proficiency with normal and emergency procedures. These techniques allowed a relatively aggressive night test program without compromising safety. A total of 18 different aircraft configurations were successfully tested, with 146 developed spins completed throughout the course of 81 program flights.

关键词： Angle of Attack flight test programme Control Surfaces Angle of Sideslip Edwards Air Force Base Conformal Fuel Tank Control Augmentation System Stabilator Crew Member Digital Electronic Engine Control

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