More effective aircraft stability and control flight testing throughuse of system identification technology

Flight simulator for civil aircraft can be used for pilot training. The flight test data for modelling, verification, validation, and qualification of flight simulator can be obtained through flight simulator data package flight test. In order to improve the efficiency of flight test and save the resources of flight test, the planning of flight simulator data package flight test was studied. Four principles of flight test planning were proposed, considering the requirements of flight simulator development progress, flight test planning of aircraft, aircraft configuration plan, requirements of flight simulator data package flight test and aircraft flight test. The principles included the principle of flight test method similarity, the principle of progress priority, the principle of pre-test meeting and the principle of considering configuration requirements. The principles of flight test planning given in this paper have been applied in the actual planning of flight simulator data package flight test for a certain aircraft, and the efficiency of flight test has been improved.

To ensure the safety of the flight control law, several development stages shall be passed, starting from the software in the loop simulators, hardware in the loop simulators, and finally the flight tests. During the flight test, the flight test data can be used to improve the aerodynamic parameter inside the flight mechanical model which is used during the hardware in the loop simulators stage. The improved aerodynamic parameters are responsible to guarantee the flight mechanical model generates the aircraft dynamics as close as possible with the real UAV at several trim points. The flight mechanical model with improved aerodynamic parameters normally called the improved flight mechanical model. The flight control law which was developed by using the improved flight mechanical model can ensure the UAV performs the auto flight as closely as possible, between the simulation and the real flight. Considering the importance of the flight test data, there is a need to develop the independent flight test instrumentation (FTI). To date, the UAV flight test data are recorded by using the flight control computer (FCC). However, there is no guarantee that the FCC record the correct data. Therefore, the independent FTI is needed for the comparative device as well as the backup device for recording the flight test data. This paper will discuss the design of a flight test instrument (FTI) system as well as a performance comparison with the FCC during the flight test. The performance comparison will compare the roll, pitch, yaw angle, and altitude (GNSS & barometer) parameters which are acquired by the FTI and the FCC. The comparison shows that the parameters obtained from the FTI have a similar trend to the FCC's. However, there is a slightly different installation location along the X-axis of the UAV, between the FCC and the FTI. That difference makes the value of pitch and yaw angle acquired by the FTI are different from the value from the FCC, while the value of the roll angle stays the same.

It is state-of-the-art to model helicopter dynamics by low-order equivalent linear systems at certain speed-dependent operating points for the flight envelope of interest. These models are obtained from flight test data by system identification. Based on the linear models, it is desired to simulate maneuvering flight such as acceleration and deceleration between hover and high forward-speed. Additionally, the identified operating point model should consist of a high modeling accuracy. Thus, the two modeling and simulation goals are: generation of a full-envelope simulation by embedding all linear models and improvement of the modeling accuracy at each operating point. These two goals stimulate the application of two techniques: model stitching and inverse simulation. Model stitching allows to implement a full-envelope quasi-nonlinear helicopter simulation whereas inverse simulation allows to analyze and model additional non-physical transfer functions that improve the linear model accuracy at a certain operating point. These transfer functions describe the remnants that are originally not considered by the system identification results. This paper presents the combination of the two techniques model stitching and inverse simulation as well as their application. Finally, the paper assesses the modeling accuracy for ACT/FHS flight test data.

The US Air Force’s two main aeropropulsion test centers, Arnold Engineering Development Center and the Air Force Flight Test Center, are developing a common suite of modeling and simulation tools employing advanced predictive modeling technologies. This common set of modeling and simulation tools incorporates real-time data validation, system identification, parameter estimation model calibration, and automated model updating as new test results or operational data become available. The expected benefit is improved efficiency and accuracy for online diagnostic monitoring of Air Force assets. These resultant models could also be used for flight manual development, determining compliance to specifications, or to aid in real-time equipment monitoring. This paper describes the integrated approach to system identification, parameter estimation, and model updating. Implementation of a software package to enable efficient model handoff between test groups and centers is discussed. An F/A-22 inlet model is used to demonstrate the approach. Compact polynomial function models of the distortion and recovery flow descriptors and 40-probe pressure values are derived from quasi-steady and instantaneous subscale wind tunnel data. The model parameters are then calibrated with F/A-22 flight test data. Results show that the modeling algorithm captures the relevant nonlinear physics of the application, and the calibration and updating procedure improves the model match to flight data. A companion paper provides preliminary results from integrating the calibrated total-pressure inlet distortion and recovery models into a real-time equipment health monitoring system to support test operations.

The present contribution describes the set-up and major results of a low speed test with an Airbus A320 wind tunnel model in high lift configuration in the ETW under cryogenic high pressure test conditions. The test is embedded in the joint LuFo IV research project HINVA (High Lift INflight Validation). The overall objective of the HINVA project is to improve the capability of CFD tools and wind tunnel testing to predict maximum lift and its determining flow phenomena in terms of accuracy and reliability. The reference for both simulation approaches is given by dedicated flight test data obtained in a flight test campaign using DLR’s ATRA test aircraft. In particular, the static aero-elastic deformations from the flight test at maximum lift conditions have been incorporated in the wind tunnel model. Moreover, attention is paid to the scaled representation of geometric features, such as slat tracks or a de-icing pipe between slat and fixed wing. The peniche height of the A320 half model is adapted to maximum lift conditions. The model is equipped with a high resolution pressure tap distribution aligned to the pressure measurement locations of the flight test aircraft. In addition to standard force, moment and pressure measurements, cryo PIV velocity field measurements have been carried out above the wing aft of the engine, as well as transitions measurements based on temperature sensitive paint. The deformation under wind loads is detected by a stereo pattern tracking system. After introducing the considerations for the model layout, the test techniques in use are introduced with selected results presented for each technique.

Nonclassical flutter has been observed during transonic conditions associated with flight and wind tunnel testing of various fighter and bomber aircraft As contrasted with classical flutter the underlying mechanism has been ascribed to a phase lag between lift variations and wing motions Based on the observations that the wing oscillations were restricted to bending mode shapes similar to the free fundamental, the phenomenon was idealized as the lift induced forced oscillation of a Cantilever beam simulating the wing. Coupling with the aerodynamics occurs through the beam equation forcing term, representing the lift, and the tangency boundary condition for the unsteady transonic small disturbance equation for the latter Strip theory has been used with a suitably revised version of LTRAN2 to supply the aerodynamic input The resulting time integration algorithm provides a procedure to determine the destabilizing effects of sweepback Computations with the method closely predict flutter frequencies observed for the HiMAT canard during wind tunnel testing, as well as for the B-l wing in flight tests in two out of three cases Qualitative behavior of the oscillation agrees for these conditions For the third case, damped oscillations are predicted rather than the amplified ones observed Reasons for the discrepancy are provided, involving the second bending mode of these oscillations in contrast to the fun damental assumed by the current approach Specific recommendations for refinements of the procedure in volving inclusion of modal coupling analysis of limit cycle aspects and three dimensional phenomena are given

A nonlinear energy-based control law was flight tested on a small, fixed-wing unmanned aircraft. This paper summarizes the selected aircraft, instrumentation system, data processing techniques, system identification methods, and the control laws that were implemented. The flight test campaign used a build up approach with increasingly complex computer generated system identification excitations and then increasingly complex control laws. This build up approach allowed the team to overcome technical challenges in a progressive manner and then finally test an advanced nonlinear control law in flight. Automated multistep, frequency sweep, and multisine excitation inputs were implemented for system identification. System identification methods were leveraged to develop linear and nonlinear aerodynamic models. A servoactuator model was developed from ground test data and data processing techniques were used to condition the flight test data for analysis. A proportional-derivative attitude commanded system was implemented and tuned using pilot comments, without the use of an a priori model. A linear quadratic regulator was tuned using the linear aerodynamic model, and refined during flight tests to improve handling qualities. A port-Hamiltonian energy-based nonlinear control law was tuned in simulation using the nonlinear model, and gains refined during flight tests to improve directional tracking and perturbation response. Implementation techniques for automated system identification maneuvers, as well as feedback control using a Pixhawk and Raspberry Pi co-computer are documented and made available by means of a publicly accessible web repository. Flight test results illustrate the utility of the experimental data collection and analysis methods for testing advanced flight control schemes.

The X-31A experimental aircraft, a joint venture between Rockwell International, USA and Deutsche Aerospace, Germany, was designed and constructed to demonstrated enhanced fighter maneuverability including operation in the post stall regime. The DLR Institute of Flight Mechanics is involved in the flight testing phase by contributing its system identification capabilities for the extraction of mathematical models and aerodynamical parameters. It is a major challenge as the hasic aircraft is highly unstable and is only flown with the control system engaged. Therefore, the paper concentrates on two associated problem areas: the identification of unstable systems and the identification of open-loop aircraft dynamics from closed-loop flight test data. Both, timeand frequencydomain identification methods were applied to determine the coefhcients for the longitudinal motion of the X-31A. The techniques are characterized and representative results are to demonstrate and discuss the suitability and constraints of the individual approaches in detail.

This paper presents a lower risk alternative to traditional single step flight programs to achieve air breathing propulsion powered hypersonic unassisted flight without the need for carrier aircraft or booster rockets. This two step approach is based on incremental flight test envelope expansion using an X-plane powered by a turbine based combined cycle (TBCC) engine with a turbine engine and a dual mode ramjet (DMRJ), an engine that can operate as a ramjet or scramjet, both burning hydrocarbon fuel. The Two Step Vehicle (TSV) X-plane would be designed to be modified and flight tested in two discrete steps. For the first step, the TSV would integrate and flight test an off-the-shelf turbine engine with a surrogate DMRJ in the TBCC flowpath. The second step would refine the DMRJ design based on the first step flight and propulsion test data, ground test the updated DMRJ design, and integrate the final DMRJ configuration with the turbine engine for a complete TBCC engine flight test. The TSV X-plane would pave the way for design and development of air breathing propulsion hypersonic air vehicles for nearterm flight demonstrations and enable future military applications such as prompt global ISR and strike and commercial applications such as two-stage-to-orbit for space access.

A method for developing and performing a rotorcraft flight simulation in a bluff body airwake is investigated and compared to flight test data. This method is broadly composed of two parts: an atmospheric boundary layer computational solution that is unsteady and time-accurate, and a UH-60 flight dynamics program that can be coupled with an airwake developed from the atmospheric boundary layer solution. The flight dynamics program is a MATLAB/SIMULINK implementation of the GENHEL model. The computational fluid dynamics airwake solution is developed to recreate the airwake present for the validation case flight tests that were used to collect the flight test data. The validation case f light test data was collected from a variety of station-keeping tasks performed in the wake of a hangar on land. Two kinds of flight simulations were performed: piloted simulations in a flight simulator facility and flight simulations performed using a pilot model. Comparisons of the power spectral density analyses of pilot control stick activity reveal that pilots in the simulated flight tests use a generally higher level of control activity energy than the pilots of the f light test validation case. Limited simulator flight test data showed sizeable differences in pilot control activity for the same simulated flight task flown by different pilots. Significant unknowns remain in comparing a simulated flight to a physical flight. The pilot model was not used to compare simulated flights to validation case flights, but was useful for comparing different flight simulation cases to each other as well as for analyzing sensitivity of flight simulations to positioning error within a simulated airwake. The pilot model flight simulations show that flight tasks performed in an airwake produced using a high-fidelity atmospheric boundary layer simulation require more pilot control input power than airwakes made using a comparable mean wind speed uniform inflow.

This paper discusses the author's experience during four flight test programs to improve the characteristics of business jet aircraft. In each case, an experimental prototype was used to evaluate aerodynamic configurations for the first time. This was invaluable because in some cases aerodynamic techniques, used to make wind tunnel models behave more like the full scale aircraft, were used for the first time to improve aircraft handling characteristics. The characteristics problems solved by simple aerodynamic modifications included wing drop (roll ofo, stable deep stall and high angle of attack divergent wing rock. The effects on high angle of attack handling of the following aerodynamic devices are discussed: strips, wing fences, vortex generators, vortilons and Delta fins. It has been my privilege, over the last several years to be involved in several characteristics development programs on business jet aircraft. These programs have been very rewarding for the following reasons; the flight test aircraft was used as a true experimental vehicle, where new ideas were tested for the first time, the test pilot was an integral part of the engineering development team, the programs led to significant improvements in aircraft handling and, therefore, safety and in each case, the end result was a simple, inexpensive aerodynamic change to the aircraft. It is not unusual for there to be differences between the characteristics predicted by wind tunnel tests and those seen during flight tests. When this occurs, some developmental flight testing is required. General aviation seems to be the only flight test area where a true experimental flight test program is economically feasible and I believe that this was critical to the success of these programs. In several cases, the devices used to modify aircraft characteristics were the same as are used to reduce scale effects to make windtunnel models behave more like the aircraft. The characteristics problems solved by simple aerodynamic modifications included wing drop (roll off), stable deep stall and high angle of attack (AOA) divergent wing rock. Several hundred aerodynamic configurations and many thousands of stalls were flown during these very intense flight test programs. As many as four configurations were flown per day. Even though these were hazardous tests, no incidents or aircraft damage occurred during the tests. The effects on high AOA handling of the following aerodynamic devices was evaluated; strips, wing fences, vortex gencrators, vortilons and Delta fins.

<div class="htmlview paragraph">A method is presented which permits determination of of propulsive efficiency and airplane drag terms from steady level flight speed-power data. The method is based on a computer formulation of Lock's equivalent propeller model which makes it suitable for use in parameter identification. The results are in the form of coefficients of polynomials defining the airplane and the propeller. These coefficients, determined from steady flight data, then permit easy calculation of propeller characteristics and airplane performance for other flight conditions. The method is particularly applicable to performance prediction and to simulation because the airplane and propeller are represented by a simple analytical model rather than by tables or graphs.</div> <div class="htmlview paragraph">The method is described, computer formulation of the propeller model is verified with wind tunnel data, airplane and propeller characteristics are determined from actual flight test data, and predicted power is compared with measured power for a given flight test condition.</div>

Many different sensor types are used in aircraft and aerospace applications in general. During the certification period of a commercial airliner, a loads survey is conducted to collect aerodynamic performance data by measuring the pressure on the surface of the wings. To do this, we developed a unique multi-MEMS pressure sensor system targeted for flight test applications. Here we describe test results collected from two aircraft flight tests using this thin profile ‘pressure belt’ multi-sensor assembly. The overall measurement accuracy was close to our targeted 0.1% of full scale, which is 10× better than the currently used technology. Lessons learned from these flight tests and approaches to further improve the pressure belt reliability are also discussed.

Methods of the actual values of the angles of attack and slip determination in the technology of air data systems estimation using satellite navigation systems during aircraft flight tests of are stated. Also methods of parametric identification of mathematical models of on-board aerodynamic angles measurements errors are presented. The technology has been developed and implemented in the practice of flight tests at Flight Research Institute. The technology shows high efficiency of solving tasks of air data on-board systems, vertical separation systems, flight performance and takeoff and landing characteristics estimation in flight tests. The technology covers a wide range of issues of aerometric measurements, many of which have already been published. Here, attention is paid to measurements of aerodynamic angles. The features of determining the angles of attack and slip using data from air data systems, satellite and inertial navigation systems are shown. Equations for determining the angles of attack and slip during testing, factors of measurement errors, the structure of the mathematical model of the aerodynamic errors of the angle of attack on-board measurements are presented in a stationary approximation. Algorithms for solving problems in steady and unsteady modes of flight, takeoff and landing are considered. The completeness of the solution for the wind speed vector is noted. Estimates of methodological errors in determining aerodynamic angles using the technology are made. Compliance with modern requirements is showing. The effectiveness of the technology is confirmed by the materials of flight tests of a short-range aircraft. Typical examples of test materials presentation, features of measurements depending on the type of test mode, aerodynamic influence factors are shown.

Comparing and benchmarking engineering methods for the prediction of X-31 aerodynamics