High Alpha Research Vehicle Research Articles

Sliding-Mode Control and Strategic Thrust-Vectoring Based Aircraft Flat Spin Recovery With Altitude Loss Minimization

Spin recovery from a fully-developed-oscillatory-stable-flat spin is performed in this article using thrust vector control (TVC) and compared with the conventional mode of recovery. TVC dynamics is modeled and then added to the standard aircraft dynamics. TVC command is designed as a function of angle of attack in the pitch and yaw direction and optimally used with aircraft primary controls (APCs). The APC is designed using the sliding-mode control technique. The spin recovery profile is demonstrated on F-18 high alpha research vehicle to test the efficacy of the proposed algorithm. Results show that with the combined strategic use of TVC and APC, spin recovery time is drastically reduced because of reduced saturation of APC. Thus, altitude decrease during spin recovery also reduces, thereby gaining on final exit altitude to level flight, compared to recovery using only APC. Thus, TVC is proven in this article to be a powerful corrective contributor to spin recovery, increasing margin of safety in terms of excessive altitude loss.

Vertical thrust-based effective flat-spin recovery of aircraft restricting altitude loss

Among the various modes of aircraft spin, flat-spin being the most ruthless form, the severity and the flight parameters govern the recovery success rate. This paper presents a novel approach for utilizing vertical thrust to reduce the fatality of the flat-spin in terms of excessive altitude loss regulating the survivability post aircraft recovery. F-18 High Alpha Research Vehicle is considered for the present study to demonstrate effectiveness of the proposed method. The dynamics of vertical thrust add-on is conceptualized and incorporated to the standard aircraft. Subsequently, to validate the effectiveness of this mode of recovery as compared to the conventional primary control based method, standard sliding-mode control technique is adopted. Closed-loop simulation shows that vertical thrust restricts altitude loss and minimizes spin recovery time compared to the conventional mode of recovery from flat-spin. The outcome of the present study ensures the use of vertical thrust as a potential solution for minimizing aircraft flat-spin-related fatal accidents.

Quantized Fault Detection Filter Design for Networked Control System with Markov Jump Parameters

In this paper, the design problem of fault detection filter for a class of networked control system subject to Markov jump parameters, randomly occurring uncertainties, sensor faults and packet dropouts by using state feedback control is investigated. The fault detection filter is used as an optimal residual generator that generates the residual signal and also guarantees the sensitivity of residual signal to the faults. To be more specific, a suitable threshold is selected for the residual evaluation function to reduce the error between the sensor faults and the residual signal. Further, by employing a fault detection technique along with output quantization approach, a new set of sufficient conditions is derived to ensure the stochastic stabilization of the addressed system with sensor faults under a predefined $$H_{\infty }$$ performance level. Particularly, by constructing Lyapunov–Krasovskii functional and utilizing extended Wirtinger-based single integral inequality, the required sufficient conditions are derived in terms of linear matrix inequalities for getting the required result. At last, two numerical examples including high alpha research vehicle model are presented to demonstrate the usefulness of our developed results.

A Codesign Methodology for Constrained Networked Systems With Frequency Specifications

This brief addresses a controller/fault detector and scheduling protocol codesign scheme in the finite frequency domain for a class of constrained networked systems. By considering the data transmission in both sensor-control center and control center-actuator channels, a switched stochastic system is first modeled. With the aid of the mode-dependent average dwell time, sufficient conditions are subsequently investigated to capture the desired performances including the finite frequency one, in which the relationship between the node accessing rate and the performances is developed. Then, node-dependent controller/fault detectors and the scheduling protocol are codesigned by the derived conditions. Finally, an application to the High Alpha Research Vehicle (HASV) is given to illustrate the effectiveness of the proposed scheme.

Nonlinear gain-scheduled flight controller design via stable manifold method

This research presents a nonlinear gain-scheduled flight controller design method via a stable manifold theory in order to handle the nonlinearities of the F-18 High Alpha Research Vehicle due to the change in the aerodynamic characteristics at different angles of attack and the airspeed variation. The designed longitudinal flight control system consists of a nonlinear gain-scheduled stabilization augmentation system which is designed using the stable manifold method, and a linear gain-scheduled control augmentation system which consists of proportional and integral gains. The nonlinear longitudinal flight controller is verified in a 6 degree-of-freedom simulator.

Modeling of unsteady aerodynamic characteristics at high angles of attack

An improved model to express the unsteady aerodynamic characteristics at high angles of attack is presented in this paper. The proposed aerodynamic model is expressed on the basis of a progressive state-space representation and Taylor’s series expansion. The state-space expression is a first-order differential equation in which the power item of the angular rate of attack is introduced. The unsteady aerodynamic coefficients are described by Taylor’s series expansion in terms of input variables. The approach of minimum mean square error criterion is utilized to identify the unknown parameters of the proposed model by nonlinear least square method from the tunnel data. The given modeling method is experimentally demonstrated by the wind tunnel measurements of NACA 0015 airfoil with constant rate to high angles of attack, F18 aircraft with constant pitch rate ramp motion, and F18 HARV (high alpha research vehicle) configuration with large-amplitude harmonic oscillatory. The results show that it is possible to analyze more complex unsteady aerodynamic problems for an aircraft within the framework of the proposed aerodynamic model and the represented model is directly amenable to the simulation and control system design.

Application of L<sub>1</sub> Adaptive Control to a High Maneuverability Aircraft

The L1 Adaptive control technique is applied to the F-18 HARV (High Alpha Research Vehicle) aircraft. This control architecture compensates for both the matched and the unmatched uncertainties of the full nonlinear aircraft dynamics. This paper describes the L1 control architecture used to control the nonlinear dynamics of the aircraft and the final part presents the detailed simulation results for the lateral channel

Networked Fault-Tolerant Control Allocation for Multiple Actuator Failures

This paper proposes intelligent fault-tolerant control technique using network. Not only control commands generated by a controller but also diagnostic data for tolerating failures can be transmitted through network. In this paper, fault-tolerant control allocation method (FTCA) is proposed to tolerate failures in more than one actuator. FTCA is based on a well-known actuator management technique called control allocation (CA). While the conventional CA is used to redistribute actuators optimally, FTCA redistributes actuators to compensate for the performance degradation due to actuator failure. To analyze the effects of faulty actuator, this paper proposes the general model of the faulty system firstly. And then the modified CA for tolerating the effect of failure is proposed. The performance of the proposed FTCA method is verified by the numerical simulations with application to F-18 High Alpha Research Vehicle (HARV).

Finite-Frequency Filter Design for Networked Control Systems with Missing Measurements

This paper is concerned with the problem of robust filter design for networked control systems (NCSs) with random missing measurements. Different from existing robust filters, the proposed one is designed in finite-frequency domain. With consideration of possible missing data, the NCSs are first modeled to Markov jump systems (MJSs). A finite-frequency stochasticH∞performance is subsequently given that extends the standardH∞performance, and then a sufficient condition guaranteeing the system to be with such a performance is derived in terms of linear matrix inequality (LMI). With the aid of this condition, a procedure of filter synthesis is proposed to deal with noises in the low-, middle-, and high-frequency domains, respectively. Finally, an example about the lateral-directional dynamic model of the NASA High Alpha Research Vehicle (HARV) is carried out to illustrate the effectiveness of the proposed method.

Aircraft Maneuver Design Using Bifurcation Analysis and Sliding Mode Control Techniques

In this paper, bifurcation analysis-based methodology is used in conjunction with a sliding-mode-based control algorithm to construct and simulate maneuvers for a nonlinear, 6 degree-of-freedom, F-18 high-alpha research vehicle aircraft model. Three different types of maneuvers, namely, a minimum-radius level turn, velocity vector roll, and spin recovery to a level flight condition, are attempted to demonstrate the usefulness of the proposed approach. The procedure involves constructing the desired maneuvers using constrained bifurcation analysis-based methodology. The results obtained from bifurcation analysis provide the reference inputs for the sliding mode controller to switch the aircraft between desired flight conditions. Robustness of the sliding mode controller is also examined by introducing uncertainties in aerodynamic parameters. Closed-loop simulation results are later presented to show the effectiveness of the proposed technique.

Susceptibility of F/A-18 Flight Controllers to the Falling-Leaf Mode: Linear Analysis

The F/A-18 Hornet aircraft with the original flight control law exhibited a nonlinear out-of-control phenomenon known as the falling-leaf mode. This unstable mode was suppressed by modifying the control law. This paper employs the falling-leaf phenomenon as an example to investigate the applicability of linear analysis tools for detecting inherently nonlinear phenomenon. A high-fidelity nonlinear model of the F/A-18 is developed for controller analysis using F/A-18 High-Alpha Research Vehicle aerodynamic data in the open literature. A variety of linear analysis methods are used to investigate the robustness properties of the original (baseline) and the revised F/A-18 flight control law at different trim points. Classical analyses, e.g., gain and phase margins, do not indicate a significant improvement in robustness properties of the revised control law over the baseline design. However, advanced robustness analyses, e.g., μ analysis, indicate that the revised control law is better able to handle the cross-coupling and variations in the dynamics than the baseline design.

Susceptibility of F/A-18 Flight Controllers to the Falling-Leaf Mode: Nonlinear Analysis

The F/A-18 Hornet aircraft with the original flight control law exhibited a nonlinear out-of-control phenomenon known as the falling-leaf mode. This unstable mode was suppressed by modifying the control law. This paper employs the falling-leaf phenomenon as an example to investigate the applicability of linear analysis tools for detecting inherently nonlinear phenomenon. A high-fidelity nonlinear model of the F/A-18 is developed for controller analysis using F/A-18 High-Alpha Research Vehicle aerodynamic data in the open literature. A variety of linear analysis methods are used to investigate the robustness properties of the original (baseline) and the revised F/A-18 flight control law at different trim points. Classical analyses, e.g., gain and phase margins, do not indicate a significant improvement in robustness properties of the revised control law over the baseline design. However, advanced robustness analyses, e.g., μ analysis, indicate that the revised control law is better able to handle the cross-coupling and variations in the dynamics than the baseline design.

On-line aircraft parameter identification using fourier transform regression with an application to NASA F/A-18 harv flight data

This paper applies a recently developed on-line parameter identification (PID) technique to sets of real flight data and compares the results with those of a state-of-the-art off-line PID technique. The on-line PID technique takes Linear Regression from Fourier Transformed equations and the off-line PID is based on the traditional Maximum Likelihood method. Sets of flight data from the NASA F/A-18 High Alpha Research Vehicle (HARV) aircraft, which has been recorded from specifically designed maneuvers and used for off line parameter estimation, are used for this study. The emphasis is given on the accuracy and on-line measure of reliability of the estimates. The comparison is performed for both longitudinal and lateral-directional dynamics for maneuvers at angles of attack ranging α=20δ through α=40δ. Results of the two estimation processes are also compared with baseline wind tunnel estimates whenever possible.

Real-Time Parameter Estimation in the Frequency Domain

A method for real-time estimation of parameters in a linear dynamic state space model was developed and studied. The application is aircraft dynamic model parameter estimation from measured data in flight for indirect adaptive or reconfigurable control. Equation error in the frequency domain was used with a recursive Fourier transform for the real-time data analysis. Linear and nonlinear simulation examples and flight test data from the F-18 High Alpha Research Vehicle HARV) were used to demonstrate that the technique produces accurate model parameter estimates with appropriate error bounds. Parameter estimates converged in less than 1 cycle of the dominant dynamic mode natural frequencies, using control surface inputs measured in flight during ordinary piloted maneuvers. The real-time parameter estimation method has low computational requirements, and could be implemented aboard an aircraft in real time.

Analysis of Random Gust Response with Nonlinear Unsteady Aerodynamics

Experimental nonlinear unsteady aerodynamics is employed to determine the maximal aircraft response to random gust through matched filter theory. Two airplane configurations, the F-18 High Alpha Research Vehicle (HARV) and F-16XL, have been tested in forced oscillation to provide the necessary data. The atmospheric turbulence is assumed to be characterized by the von Karman gust power spectral density function. The nonlinear aerodynamic models are set up through Fourier functional analysis of the test data. For comparative purposes, linear aerodynamic models are also calculated by using the unsteady quasi-vortex-lattice method. It is shown that nonlinear unsteady aerodynamics produce the instantaneous maximum lift response to gust significantly higher than that by the linear unsteady aerodynamics

Flight Test of Optimal Inputs and Comparison with Conventional Inputs

A technique for designing optimal inputs for aerodynamic parameter estimation was flight tested on the F-18 High Alpha Research Vehicle. Model parameter accuracies calculated from flight-test data were compared on an equal basis for optimal input designs and conventional inputs at the same flight condition. In spite of errors in the a priori input design models and distortions of the input forms by the feedback control system, analysis of data generated by the optimal inputs revealed lower estimated parameter errors compared with conventional 3-2-1-1 and doublet inputs. In addition, the tests using optimal input designs demonstrated enhanced design flexibility, allowing the optimal input design technique to use a larger input amplitude to achieve further increases in estimated parameter accuracy without departing from the desired flight-test condition. This work validated the analysis used to develop the optimal input designs, and demonstrated the feasibility and effectiveness of the optimal input design technique.

Wavelet-Processed Flight Data for Robust Aeroservoelastic Stability Margins

Waveletanalysisfore lteringandsystem identie cationisusedto improvetheestimationofaeroservoelastic (ASE) stabilitymargins.Computationofrobuststabilitymarginsforstabilityboundarypredictiondependsonuncertainty descriptions derived from the test data for model validation. Nonideal test conditions, data acquisition errors, and signal processing algorithms cause uncertainty descriptions to be intrinsically conservative. The conservatism of the robust stability margins is reduced with parametric and nonparametric time-frequency analysis of e ight data in the model validation process. Nonparametric wavelet processing of data is used to reduce the effects of external disturbances and unmodeled dynamics. Parametric estimates of modal stability are also extracted using the wavelet transform. F-18 High Alpha Research Vehicle ASE e ight test data are used to demonstrate improved robust stability prediction by extension of the stability boundary from within the e ight envelope to conditions sufe cently beyond the actual e ight regime. Stability within the e ight envelope is cone rmed by e ight test. Practical aspects and guidelines for efe ciency of these procedures are presented for on-line implementation.

Non-linear flight dynamics at high angles-of-attack

AbstractThe methods of non-linear dynamics are applied to the longitudinal motion of a vectored thrust aircraft, in particular the behaviour at high angles-of-attack. The model contains analytic non-linear aerodynamic coefficients based on Nasa windtunnel tests on the F-18 high alpha research vehicle (Harv). The equilibrium surfaces are plotted against thrust magnitude and thrust deflection and are used to explain the behaviour. When the aircraft is forced with small thrust deflections whilst in poststall equilibrium, chaotic motion is observed at certain frequencies. At other frequencies, several limiting states coexist, e.g. a chaotic attractor and a limit cycle, or two limit cycles. The steady state behaviour then depends on the initial conditions. The non-linear pitching moments are shown to have a significant effect upon the model. Omitting these terms, the region of chaos shrinks, this is not observed for lift.

Accuracy of Aerodynamic Model Parameters Estimated from Flight Test Data

An important put of building mathematical models based on measured date is calculating the accuracy associated with statistical estimates of the model parameters. Indeed, without some idea of this accuracy, the parameter estimates themselves have limited value. An expression is developed for computing quantitatively correct parameter accuracy measures for maximum likelihood parameter estimates when the output residuals are colored. This result is important because experience in analyzing flight test data reveals that the output residuals from maximum likelihood estimation are almost always colored. The calculations involved can be appended to conventional maximum likelihood estimation algorithms. Monte Carlo simulation runs were used to show that parameter accuracy measures from the new technique accurately reflect the quality of the parameter estimates from maximum likelihood estimation without the need for correction factors or frequency domain analysis of the output residuals. The technique was applied to flight test data from repeated maneuvers flown on the F-18 High Alpha Research Vehicle. As in the simulated cases, parameter accuracy measures from the new technique were in agreement with the scatter in the parameter estimates from repeated maneuvers, whereas conventional parameter accuracy measures were optimistic.

Comparison of F/A-18A inlet flow analyses with flight data. I

In this article NPARC calculations of the F/A-18A High Alpha Research Vehicle installed inlet flow are compared with flight test at a Mach number of 0.3, angles of attack of 30 and 60 deg, and nonzero sideslip angle. Predicted forebody, leading-edge extension, inlet lip, and duct surface pressures agree well with flight data. An examination of the flight-test database total pressures shows a significant degree of flow unsteadiness at the engine face. The computational results represent an asymptotic solution driven by steady boundary conditions and do not represent any particular point in time in the flight test. Therefore, comparisons of calculated inlet performance are made within a band of minimum and maximum flighttest values of inlet performance parameters. Comparisons of the predicted total pressure contours at the engine face with flight data indicate that the numerical calculations of inlet flow distortions and total pressure recovery at the engine face are at the upper and lower range, respectively, of test data, and the calculated total pressure contours are within the excursion range of the minimum total pressure, or pressure well, shown by flight data.