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Low-speed free flight tunnel testing of swept wing fighter aircraft for determining effect of small canard surfaces on directional stability
Pretest predictions of the dynamic characteristics of the XV-15 tilting proprotor aircraft are presented. The data for the aircraft in flight include: trim conditions, flight dynamics, gust response, aeroelastic stability and the wing response to control. The data for the aircraft in the Ames 40- by 80-ft wind tunnel include aeroelastic stability and the wing response to control. The calculations were made for pylon tilt angles of 0 deg (airplane mode), 0.1 deg (pylon unlocked), 30, 60, and 90 deg (helicopter mode).
Feasibility of boron-epoxy reinforced aluminum alloy for horizontal stabilizer structure of DC 8 aircraft and recommendations for demonstration program
A method of parameter extraction for stability and control derivatives of aircraft from flight test data, implementing maximum likelihood estimation, was developed and successfully applied to actual longitudinal flight test data from a modern sophisticated jet fighter. The results of this application establish the merits of the estimation technique and its computer implementation(allowing full analyst interaction with the program) as well as provide data for the validation of a portion of the differential maneuvering simulator (DMS). The results are presented for all flight test runs in tabular form and as time history comparisons between the estimated states and the actual flight test data. Comparisons between extracted and manufacturer's values for five major derivatives are presented and reveal good agreement for these principal derivatives with one exception. This particular derivative is extensively investigated by utilizing the interactive capabilities of the computer program. The results of this investigation verify the numbers extracted by maximum likelihood estimation.
A maximum likelihood parameter estimation procedure and program were developed for the extraction of the stability and control derivatives of aircraft from flight test data. Nonlinear six-degree-of-freedom equations describing aircraft dynamics were used to derive sensitivity equations for quasilinearization. The maximum likelihood function with quasilinearization was used to derive the parameter change equations, the covariance matrices for the parameters and measurement noise, and the performance index function. The maximum likelihood estimator was mechanized into an iterative estimation procedure utilizing a real time digital computer and graphic display system. This program was developed for 8 measured state variables and 40 parameters. Test cases were conducted with simulated data for validation of the estimation procedure and program. The program was applied to a V/STOL tilt wing aircraft, a military fighter airplane, and a light single engine airplane. The particular nonlinear equations of motion, derivation of the sensitivity equations, addition of accelerations into the algorithm, operational features of the real time digital system, and test cases are described.
A method of parameter extraction for stability and control derivatives of aircraft from flight test data, implementing maximum likelihood estimation, has been developed and successfully applied to actual lateral flight test data from a modern sophisticated jet fighter. This application demonstrates the important role played by the analyst in combining engineering judgment and estimator statistics to yield meaningful results. During the analysis, the problems of uniqueness of the extracted set of parameters and of longitudinal coupling effects were encountered and resolved. The results for all flight runs are presented in tabular form and as time history comparisons between the estimated states and the actual flight test data.
A conceptual design study was conducted to define a representative military and/or commercial tilt-propeller aircraft for short takeoff and landing operation. The level of structural technology selected for the operational aircraft was based on aluminum, steel, titanium, and adhesive bonded structures. The data describe the following: (1) aircraft weight, (2) performance and stability, (3) aerodynamic noise, (4) dynamic characteristics, (5) maintainability and reliability, and (6) operating economics.
A maximum likelihood estimation method was applied to flight data and procedures to facilitate the routine analysis of a large amount of flight data were described. Techniques that can be used to obtain stability and control derivatives from aircraft maneuvers that are less than ideal for this purpose are described. The techniques involve detecting and correcting the effects of dependent or nearly dependent variables, structural vibration, data drift, inadequate instrumentation, and difficulties with the data acquisition system and the mathematical model. The use of uncertainty levels and multiple maneuver analysis also proved to be useful in improving the quality of the estimated coefficients. The procedures used for editing the data and for overall analysis are also discussed.
A stoppable helicopter rotor with circulation control was investigated in the Ames 40 by 80 foot wind tunnel. The model was tested as a rotating wing, a fixed wing, and during transition start/stop sequences. The capability of the model's control system to maintain pitch and roll moment balance during the start/stop sequence, the ability of the blades to withstand the start/stop loads, the adequacy of the control system to maintain balance in the helicopter mode, and the control system capabilities in the fixed-wind mode were assessed. Time-history data of several start/stop sequences of the X-wing rotor, and the steady-state data relating to the model as both a rotor and as a fixed-wing aircraft are presented. In addition, stability data are presented which were acquired during open-loop and closed-loop tests of the hub moment feedback control system.
NASA's F-8 digital fly-by-wire aircraft has been subjected to stability and control derivative assessments, leading to the proposal of improved control inputs for more efficient control derivative estimation. This will reduce program costs by reducing flight test and data analysis requirements. Inputs were divided into sinusoidal types and cornered types. Those with corners produced the best set of stability and control derivatives for the unaugmented flight control system mode. Small inputs are noted to have provided worse derivatives than larger ones.
Flight test results are presented documenting the effect of airframe icing on performance and stability and control of a NASA DHC-6 icing research aircraft. Kohlman System Research, Inc., provided the data acquisition system and data analysis under contract to NASA. Performance modeling methods and MMLE techniques were used to determine the effects of natural ice on the aircraft. Results showed that ice had a significant effect on the drag coefficient of the aircraft and a modest effect on the MMLE derived longitudinal stability coefficients (code version MMLE). Data is also presented on asymmetric power sign slip maneuvers showing rudder floating characteristics with and without ice on the vertical stabilizer.
An important step in system identification of aircraft is the estimation of stability and control derivatives from flight data along with an assessment of parameter accuracy. When the maximum likelihood estimation technique is used, parameter accuracy is commonly assessed by the Cramer-Rao lower bound. It is known, however, that in some cases the lower bound can be substantially different from the parameter variance. Under these circumstances the Cramer-Rao bounds may be misleading as an accuracy measure. This paper discusses the confidence interval estimation problem based on likelihood ratios, which offers a more general estimate of the error bounds. Four approaches are considered for computing confidence intervals of maximum likelihood parameter estimates. Each approach is applied to real flight data and compared.
The Shuttle Training Aircraft (STA) is a variable stability, variable control law flying simulator used by NASA/JSC to train astronauts in the final landing phase of a Space Shuttle Orbiter. A general outline is given for the STA flight simulation system. An overview is given of the software generation and verification process through the Advanced Validation System (AVAS). The flight test techniques for software verification will be reviewed and the process for releasing the software for flight training will be covered. The astronaut STA training syllabus is examined. Parameter matching with the Orbiter in the final approach phase of de-orbit and landing is briefly examined. Simulation performance will be assessed against flight data, performance measurement, and cue synchronization.
The most promising method of achieving significantly lower skin-friction drag of an aircraft component is through the stabilization and maintenance of a laminar boundary layer over a large fraction of its surface. In a number of applications, however, laminar flow is unfeasible and/or undesirable, and the most efficient turbulent flow design must be provided. The development of both laminar and turbulent airfoil and wing designs over the past decade is reviewed. High, medium, and low-speed airfoil concepts are discussed, and some overall comparisons are made for different Reynolds number ranges.
A methodology is presented for a modal suppression control law design using flight test data instead of mathematical models to obtain the required gain and phase information about the flexible airplane. This approach is referred to as BODEDIRECT. The purpose of the BODEDIRECT program is to provide a method of analyzing the modal phase relationships measured directly from the airplane. These measurements can be achieved with a frequency sweep at the control surface input while measuring the outputs of interest. The measured Bode-models can be used directly for analysis in the frequency domain, and for control law design. Besides providing a more accurate representation for the system inputs and outputs of interest, this method is quick and relatively inexpensive. To date, the BODEDIRECT program has been tested and verified for computational integrity. Its capabilities include calculation of series, parallel and loop closure connections between Bode-model representations. System PSD, together with gain and phase margins of stability may be calculated for successive loop closures of multi-input/multi-output systems. Current plans include extensive flight testing to obtain a Bode-model representation of a commercial aircraft for design of a structural stability augmentation system.
The study addresses the present transition research directed to the prediction of transition on the wings of modern transonic aircraft and concentrates on the stability of the types of boundary layers that arise on real swept wings in compressible flow. The need to perform experiments to get better insights into such aspects of transition as receptivity, types of nonlinear interaction, development of irregular chaotic motion, and the final breakdown stage is emphasized. The use of low-speed wind tunnel is discussed, and focus is placed on control, surface quality, and instrumentation.
This document is a preliminary design of a High Speed Civil Transport (HSCT) named the RTJ-303. It is a 300 passenger, Mach 1.6 transport with a range of 5000 nautical miles. It features four mixed-flow turbofan engines, variable geometry oblique wing, with conventional tail-aft control surfaces. The preliminary cost analysis for a production of 300 aircraft shows that flyaway cost would be 183 million dollars (1992) per aircraft. The aircraft uses standard jet fuel and requires no special materials to handle aerodynamic heating in flight because the stagnation temperatures are approximately 130 degrees Fahrenheit in the supersonic cruise condition. It should be stressed that this aircraft could be built with today's technology and does not rely on vague and uncertain assumptions of technology advances. Included in this report are sections discussing the details of the preliminary design sequence including the mission to be performed, operational and performance constraints, the aircraft configuration and the tradeoffs of the final choice, wing design, a detailed fuselage design, empennage design, sizing of tail geometry, and selection of control surfaces, a discussion on propulsion system/inlet choice and their position on the aircraft, landing gear design including a look at tire selection, tip-over criterion, pavement loading, and retraction kinematics, structures design including load determination, and materials selection, aircraft performance, a look at stability and handling qualities, systems layout including location of key components, operations requirements maintenance characteristics, a preliminary cost analysis, and conclusions made regarding the design, and recommendations for further study.
This Symposium is intended to bring together the often distinct cultures of the Stability and Control (S&C) community and the Computational Fluid Dynamics (CFD) community. The COMSAC program is itself a new effort by NASA Langley to accelerate the application of high end CFD methodologies to the demanding job of predicting stability and control characteristics of aircraft. This talk is intended to set the stage for needing a program like COMSAC. It is not intended to give details of the program itself. The topics include: 1) S&C Challenges; 2) Aero prediction methodology; 3) CFD applications; 4) NASA COMSAC planning; 5) Objectives of symposium; and 6) Closing remarks.