Modeling and Computing Worst-Case Uncertainty Combinations for Flight Control Systems Analysis

WORST-CASE UNCERTAIN PARAMETER COMBINATIONS FOR FLIGHT CONTROL SYSTEMS ANALYSIS

Reliability of flight control system (FCS) is a key factor of the aircraft safety. Therefore, redundant technology and Built-in Test (BIT) technology are usually adapted by flight control system for modern large civil aircraft so as to improve the system reliability. Taking Boeing 777 as an example, Imperfect Fault Coverage Model (IFCM) is built with the analysis on the influence of dissimilarity technology, redundant bus, and false alarm of BIT technology. Finally, the comparisons on reliability calculation between the IFCM proposed in this paper and other analytical models are presented. The computational result shows that, by taking into account more complex conditions, the proposed method is of great value in engineering applications and analysis.

PurposeThe paper aims to provide further study on the development and analysis of flight control system for two‐dimensional (2D) differential geometric (DG) guidance and control system based on the application of a set‐point weighting proportional‐integral‐derivative (PID) controller.Design/methodology/approachThe commanded angle‐of‐attack is developed in the time domain using the classical differential geometry theory. Then, a set‐point weighting PID controller is introduced to develop a flight control system so as to form the 2D DG guidance and control system, and the gains of the PID controller are determined by the Ziegler‐Nichols method as well as the Routh‐Hurwitz stability criterion. Finally, the classical frequency method is utilized to study the relative stability and robustness of the designed flight control system.FindingsThe results demonstrate that the designed controller yields a fast responding and stable system which is robust to the high frequency parameters variation. Moreover, the DG guidance law is viable and effective in a realistic missile defense engagement.Originality/valueThis paper provides a novel approach on the development of DG guidance and control system associated with its stability analysis.

This paper presents an integrated flight dynamic modeling method for flexible aircraft that captures coupled physics effects due to inertial forces, aeroelasticity, and propulsive forces that are normally present in flight. The present approach formulates the coupled flight dynamics using a structural dynamic modeling method that describes the elasticity of a flexible, twisted, swept wing using an equivalent beam-rod model. The structural dynamic model allows for three types of wing elastic motion: flapwise bending, chordwise bending, and torsion. Inertial force coupling with the wing elasticity is formulated to account for aircraft acceleration. The structural deflections create an effective aeroelastic angle of attack that affects the rigid-body motion of flexible aircraft. The aeroelastic effect contributes to aerodynamic damping forces that can influence aerodynamic stability. For wing-mounted engines, wing flexibility can cause the propulsive forces and moments to couple with the wing elastic motion. The integrated flight dynamics for a flexible aircraft are formulated by including generalized coordinate variables associated with the aeroelastic-propulsive forces and moments in the standard state-space form for six degree-of-freedom flight dynamics. A computational structural model for a generic transport aircraft has been created. The eigenvalue analysis is performed to compute aeroelastic frequencies and aerodynamic damping. The results will be used to construct an integrated flight dynamic model of a flexible generic transport aircraft.

State estimation for semi-Markovian switching CVNNs with quantization effects and linear fractional uncertainties

The developing field of probabilistic design has matured to the point where several classes of analysis methods have been proven to be useful for engineering analysis of large high fidelity structures with uncertain parameters. However, several barriers still stand in the way of widespread acceptance of probabilistic methods into the design practice. These include lack of data for characterizing probabilistic inputs, lack of understanding of analysis methodologies and associated limitations, lack of commercially available tools for performing probabilistic analysis, and lack of verification and validation examples. As a first step towards developing more computationally robust tools for performing probabilistic analysis and improving user understanding of analysis assumptions and limitations, this paper identifies the various types of error that can be encountered when performing probabilistic analysis and offers some suggestions for mitigating or eliminating them from the analysis.

During ground manoeuvres, particularly at high velocity when exiting the runway, a loss of lateral stability due to the saturation of the main landing gear tyres can cause the aircraft to enter a skid or a spin. The lateral stability is governed not only by aspects of the gear design, such as its geometry and tyre characteristics but also by operational variables, for example, the weather and taxiway condition. In this paper we develop an improved understanding and a new presentation of the dynamics of an aircraft manoeuvring on the ground, ultimately aimed at optimisation and automation of ground operations. Using a modified version of a nonlinear computer model of an A320 passenger aircraft developed by the Landing Gear Group at Airbus in the UK, we apply numerical continuation to investigate turning manoeuvres. We present a bifurcation analysis to describe the underlying structure that governs the dynamics of turning manoeuvres across the steering angle-thrust parameter space. There is a detailed description of the behaviour when lateral stability is lost using a novel diagrammatic approach to explain the dynamic tyre ground interactions. The bifurcation diagrams identify parameter regions for which this undesirable behaviour is avoidable and, thus, they form a foundation for defining the safe operating limits during turning manoeuvres.

This paper describes a dual-loop control scheme for fault tolerant flight control system design. The dual-loop controller consists of an outer loop controller–so-called adaptive neural sliding mode control (ANSC) and an inner loop controller designed by using nonlinear dynamic inversion (NDI) technique. The merits of adaptive neural network and sliding mode control scheme are that 1) the ability of adaptive neural network control to deal with unstructured uncertainty and 2) the ability of sliding mode control to guarantee transient response. Using timescale separation principal, the aircraft dynamics can be decomposed into fast and slow dynamics and the decomposed dynamics are inversed for NDI controllers. For real-time pilot simulation, one-stage inverse dynamics is used and the pilot inputs are translated to roll, pitch and yaw rate commands. For cascade NDI, two-stage dynamic inversion is used. The stability analysis of the proposed controller is performed using Lyapunov theory. To verify the effectiveness of the proposed control scheme, numerical simulation is performed for six degree-of-freedom nonlinear aircraft model while a failure occurs in longitudinal control surface. Simulation results demonstrate that closedloop system has good performance while encountering lock-in-place, partial destruction and floating actuator failures. Nomenclature stick long δ , stick lat δ , stick dir δ = pilot inputs for longitudinal, lateral and directional command dir lon lat K K K , , = sticks and pedal gains ref p , ref q , ref r = reference model rate commands cmd z n , cmd y n = normal and lateral acceleration command

The harmonic balance method is a very useful tool for characterizing and predicting the response of nonlinear dynamic systems undergoing periodic oscillations, either self-excited or due to harmonic excitation. The method and several of its variants were applied to nonlinear aeroelastic systems over the last two decades. This paper presents a detailed description of several harmonic balance methods and a continuation framework allowing the methods to follow the response of dynamic systems from the bifurcation point to any desired parameter value, while successfully negotiating further fold bifurcations. The continuation framework is described for systems undergoing subcritical and supercritical Hopf bifurcations as well as a particular type of explosive bifurcation. The methods investigated in this work are applied to a nonlinear aeroelastic model of a generic transport aircraft featuring polynomial or free-play stiffness nonlinearity in the control surface. It is shown that high-order harmonic balance solutions will accurately capture the complete bifurcation behavior of this system for both types of nonlinearity. Low-order solutions can become inaccurate in the presence of numerous folds in the limit-cycle oscillation branch but can still yield practical engineering information at a fraction of the cost of higher-order solutions. Time-domain harmonic balance schemes are shown to be more computationally expensive than the standard harmonic balance approach.

This paper studies the attitude control problem for the longitudinal motion of a hypersonic aircraft model containing 28 inertial and aerodynamic uncertain parameters given by NASA Langley Research Center. To handle the nonlinear dynamics and uncertain parameters, ADRC for the unobservable and unmatched disturbance is designed. Besides, the properties of the closed-loop system is discussed in two aspects: stability and robustness to the uncertainties. The stability of the closed-loop system is proved despite of nonlinear uncertain dynamics. Besides, this paper points out that the closed-loop control system is sensitive to five uncertain parameters and it also gives the stability regions of these five uncertain parameters. Finally, numerical simulation exercises are made to validate the theoretical analysis.

Experimental evaluations have been conducted of blown high-lift devices and control surfaces applied to improve the low-speed terminal-area performance of generic High Speed Civil Transport (HSCT) aircraft. Plain blown flaps and advanced pneumatic high-lift devices have been integrated into highly swept vortex-dominated wings. These produced large lift increases, as well as significant drag reductions greater than full jet thrust recovery. Because conventional horizontal tails were inadequate to trim these configurations, blown canards were employed. In addition to providing positive lift increments for pitch trim, the canards were found to favorably influence the higher angle-of-attack vortex-lift characteristics of the wings. The downwash from these canards resulted in a delay of wing vortex burst. Details of these investigations and test results that confirm the effectiveness of combined pneumatic high-lift devices and control surfaces on these HSCT aircraft configurations are presented

Let K be an arbitrary compact convex set of square matrices. This paper presents necessary and sufficient conditions for the spectrum of K to have no eigenvalues in a prespecified closed convex subset of the complex plane. The obtained result implies different criteria for analysis of the spectral set of K. In particular, we have formulated criteria for nonsingularity, inertia and Hurwitz stability of K which can be used in the robustness analysis of linear control systems with uncertain parameters.

In this paper, we first clarify the difference between the approximate z transform and the discrete equivalent of a continuous system using higher-order integrators. It is shown that a 1/ ts factor needs to be included for the approximate z transform but not for the discrete equivalent. We further apply the approximate z transform to facilitate the stability analysis of sampled-data control systems, with or without uncertain parameters, ft is shown in this paper that the approximate z transform greatly simplifies the stability analysis of a sampled-data control system, which is regarded as rather difficult ( if not impossible) to handle because of its transcendental nature. The results can be easily obtained and show reasonably good approximations with this approach. Several examples are used to illustrate the effectiveness of this new method.

The aim of this work is to present a controller based on the use of a neural network for the gust alleviation. Such alleviator acts on the lon- gitudinal control producing the necessary variation law of the elevator angle to reduce the effects of a simmetrical gust. The typology of the network, its architecture, the training procedure,as well as the controller structure, have been optimised and are herein presented. The proposed method is useful as an off-line con- trol technique, as a matter of fact first the aircraft dynamics is identified and secondly a controller is designed for it. A case study involving a typical aircraft, whose dynamics is assumed to be non-linear, shows a good performance of this controller. As a matter of fact the gust cases considered, schematized by wind speed components having sinusoidal, step and sto- castic trends, properly cover the possible case record and represent a a good test bench for the controller. Results are presented as comparisons between the performances over time of the desired state vari- ables and the actual state variables subsequent to the real aircraft dynamics under (subsequent to) the control input.

Results are presented for the initial flight tests of a novel in-line retrofit reconfigurable control technique. The technique is designed as an add-on module that is intended to upgrade rather than replace the existing control laws of current aircraft. The technique achieves reconfiguration by augmenting the pilot inceptor commands rather than the actuator commands. That is, the in-line method augments the inputs rather than the outputs of the existing control law. This architecture was chosen to balance reconfiguration performance with expected software certification requirements. The retrofit reconfigurable control algorithms were hosted on specialized research flight control computers that were installed in an F/A-18C for flight testing. Handling qualities were evaluated for simulated aerodynamic control surface failures during four test flights. The flight test results show that the technique is effective in restoring acceptable flying qualities following a change in the aircraft dynamics. In ninety-five percent of the cases tested, the system with the retrofit reconfigurable controller performed as good as, or better than, the baseline system. The most significant improvements often result from the more aggressive maneuvers with the most severe control surface failures. Verification and validation of the in-line architecture requires further study to determine to what degree the certification requirements may be lessened as compared with those of other retrofit architectures.