Nonlinear Flight Control System Research Articles

Design of sliding mode flight control system for a flexible aircraft

The evolution of large transport aircraft is characterized by longer fuselages and larger wingspans, while efforts to decrease the structural weight reduce the structural stiffness. Both effects lead to more flexible aircraft structures with significant aeroelastic coupling between flight mechanics and structural dynamics, especially at high speed, high altitude cruise. The lesser frequency separation between rigid body and flexible modes of flexible aircraft results in a stronger interaction between the flight control system and its structural modes, with higher flexibility effects on aircraft dynamics. Therefore, the design of a flight control law based on the assumption that the aircraft dynamics are rigid is no longer valid for the flexible aircraft. This paper focuses on the design of a flight control system for flexible aircraft described in terms of a rigid body mode and four flexible body modes and whose parameters are assumed to be varying. In this paper, a conditional integral based sliding mode control (SMC) is used for robust tracking control of the pitch angle of the flexible aircraft. The performance of the proposed nonlinear flight control system has been shown through the numerical simulations of the flexible aircraft. Good transient and steady-state performance of a control system are also ensured without suffering from the drawback of control chattering in SMC.

A Novel Sliding Mode Controller for Underactuated Vertical Takeoff and Landing Aircraft

Compared with other control methods, the biggest advantage of using sliding mode variable structure control method lies in its strong robustness which could be used to directly handle the strong nonlinear flight control system. However, this control method requires switching between different switching surfaces, which will inevitably cause buffeting problems, so that the energy consumption increases. Therefore, how to overcome this disadvantage to achieve the superior performance of sliding mode variable structure control method is the current research focus. This paper studies the trajectory tracking of under-actuated VTOL aircraft with three degrees of freedom and two control inputs under various coupling effects. By the input and coordinate transformation, the dynamic equation of the system is transformed into decoupled standard under-actuated form and the sliding mode controller is designed. Then Lyapunov stability theorem is used to derive sliding mode control law which could ensure that the system asymptotically converges to the given trajectory. The simulation has demonstrated the effectiveness of this method

Robustness Analysis of Modified Incremental Nonlinear Dynamic Inversion for Small UAVs

The work presented in this research focuses on the design of a robust nonlinear flight control system for a small fixed-wing UAV against uncertainties and external disturbances. Toward this objective, an integrated UAV waypoints guidance scheme based on Carrot Chasing guidance law (CC) is adopted. The designed attitude angles guidance law is applied to the flight control loop. Nonlinear Dynamic Inversion (NDI) awards the flight control system researchers a straight forward method of deriving control laws for nonlinear systems. The control inputs are used to eliminate unwanted terms in the equations of motion using negative feedback of these terms. However precise dynamic models may not be available and that leads to system instability in the presence of external disturbances or model uncertainties. Therefore a modified incremental dynamic inversion (MINDI) is presented to compensate the model uncertainties and increase robustness and system behavior when compared to incremental NDI. Simulation results showed that the MINDI flight control system is robust against wind disturbances and model mismatch which leads to superior path following performance.

Design and Analysis for Early Warning of Rotor UAV Based on Data-Driven DBN

The unmanned aerial vehicle (UAV), which is a typical multi-sensor closed-loop flight control system, has the properties of multivariable, time-varying, strong coupling, and nonlinearity. Therefore, it is very difficult to obtain an accurate mathematical diagnostic model based on the traditional model-based method; this paper proposes a UAV sensor diagnostic method based on data-driven methods, which greatly improves the reliability of the rotor UAV nonlinear flight control system and achieves early warning. In order to realize the rapid on-line fault detection of the rotor UAV flight system and solve the problems of over-fitting, limited generalization, and long training time in the traditional shallow neural network for sensor fault diagnosis, a comprehensive fault diagnosis method based on deep belief network (DBN) is proposed. Using the DBN to replace the shallow neural network, a large amount of off-line historical sample data obtained from the rotor UAV are trained to obtain the optimal DBN network parameters and complete the on-line intelligent diagnosis to achieve the goal of early warning as possible as quickly. In the end, the two common faults of the UAV sensor, namely the stuck fault and the constant deviation fault, are simulated and compared with the back propagation (BP) neural network model represented by the shallow neural network to verify the effectiveness of the proposed method in the paper.

Design and Analysis of a Robust UAV Flight Guidance and Control System Based on a Modified Nonlinear Dynamic Inversion

The work presented in this paper focuses on the design of a robust nonlinear flight control system for a small fixed-wing UAV against uncertainties and external disturbances. Toward this objective, an integrated UAV waypoints guidance scheme based on Carrot Chasing guidance law (CC) in comparison with the pure pursuit and line of sight-based path following (PLOS) guidance law is analyzed. For path following based on CC, a Virtual Track Point (VTP) is introduced on the path to let the UAV chase the path. For PLOS, the pure pursuit guidance law directs the UAV to the next waypoint, while the LOS guidance law steers the vehicle toward the line of sight (LOS). Nonlinear Dynamic Inversion (NLDI) awards the flight control system researchers a straight forward method of deriving control laws for nonlinear systems. The control inputs are used to eliminate unwanted terms in the equations of motion using negative feedback of these terms. The two-time scale assumption is adopted here to separate the fast dynamics—three angular rates of aircraft—from the slow dynamics—the angle of attack, sideslip, and bank angles. However, precise dynamic models may not be available, therefore a modification of NLDI is presented to compensate the model uncertainties. Simulation results show that the modified NLDI flight control system is robust against wind disturbances and model mismatch. PLOS path-following technique more accurately follows the desired path than CC and also requires the least control effort.

Examples on increased-order aeroservoelastic modeling

Aeroservoelastic loads calculations usually rely on linear analyses that use implicit structural models (Finite Elements Method) and frequency domain unsteady aerodynamics (Doublet Lattice Method). The hypotheses that underpin this methodology impose some limitations that may be overcome in some cases of interest. This paper details some of these cases, examples in which Increase Order Modeling has been used in aeroservoelastic and structural dynamic problems to include effects that are generally not taken into account in standard linear analyses. Increase order modeling is based on the hypothesis that the problem is mainly linear and that non-linear effects are known and concentrated. The linear part is solved in the frequency domain and convolution integrals are used to account for non-linear effects. During the last years, this methodology has been included as the standard tool at the Structural Dynamics and Aeroelasticity department of Airbus Defence and Space, allowing for tackling different problems. The first type of problems are linked to overcoming limitations imposed by the unsteady aerodynamics hypotheses; these include: taking into account aerodynamic forces dependent of in-plane motions, accounting for in-plane aerodynamic forces and considering the change in direction of aerodynamic forces during the analyses. The second group take into account non-linear effects, as the inclusion of non-linear Flight Control System for gust response calculation. Also structural non-linearities, such as actuator freeplay, contacts that change boundary conditions, and the rupture of structural elements have been implemented in dynamic problems in which frequency domain unsteady aerodynamic is considered. The Increased Order Modeling tool has been designed to fit seamlessly with the standard aeroservoelastic loads calculation methodology and models.

PI(D) tuning for Flight Control Systems via Incremental Nonlinear Dynamic Inversion

Previous results reported in the robotics literature show the relationship between time-delay control (TDC) and proportional-integral-derivative control (PID). In this paper, we show that incremental nonlinear dynamic inversion (INDI) — more familiar in the aerospace community — are in fact equivalent to TDC. This leads to a meaningful and systematic method for PI(D)-control tuning of robust nonlinear flight control systems via INDI. We considered a reformulation of the plant dynamics inversion which removes effector blending models from the resulting control law, resulting in robust model-free control laws like PI(D)-control.

Design of the Nonlinear Structure Adaptive Model Inversion Flight Control System

[abstFig src='/00280006/17.jpg' width='300' text='Design of the nonlinear SAMI flight control system' ] The advanced aircraft has those technological characteristics: stealth, supersonic cruise, super maneuverability, multi-target attack, multi-role, high load, long range cruise, integrated avionics, short takeoffs and vertical landings. These characteristics need multi-subsystem integration in the advanced aircraft. In order to meet the need of advance aircraft, the nonlinear control system is necessary. In this paper, the nonlinear structure adaptive model inversion system was first used for the nonlinear control problem of super-maneuver aircraft. The sufficient and necessary condition of the control law was analyzed. Finally the results of the height-angle maneuver simulation show that the designed control system has good performance.

Fault Tolerant Controller Design for a Faulty UAV Using Fuzzy Modeling Approach

We address a fault tolerant control (FTC) issue about an unmanned aerial vehicle (UAV) under possible simultaneous actuator saturation and faults occurrence. Firstly, the Takagi-Sugeno fuzzy models representing nonlinear flight control systems (FCS) for an UAV with unknown disturbances and actuator saturation are established. Then, a normal H-infinity tracking controller is presented using an online estimator, which is introduced to weaken the saturation effect. Based on the normal tracking controller, we propose an adaptive fault tolerant tracking controller (FTTC) to solve actuator loss of effectiveness (LOE) fault problem. Compared with previous work, this approach developed in our research need not rely on any fault diagnosis unit and is easily applied in engineering. Finally, these results in simulation indicate the efficiency of our presented FTC scheme.

Nonlinear Flight Control System Design Including Handling Qualities Evaluation

Progress in aircraft designs heavily depend on automatic flight control system. Modern aircraft have a wide flight envelope (linear and nonlinear) region. This study considers the design of stability and control augmentation system (SCAS) in high angle of attack region for a modern type aircraft via back stepping method. Because of highly nonlinear phenomena in aerodynamic parameters, a locally valid linear in the parameters nonlinear aircraft model is defined, of which the unknown parameters are approximated by Lyapunov base update laws. The numerical simulations show that the controller is capable of control aircraft based on conventional handling qualities and has a better time response than the classical PI.

Linear Aerodynamic Model Identification of a Flapping Wing MAV Based on Flight Test Data

This paper presents an approach to the system identification of the Delfly II Flapping Wing Micro Air Vehicle (FWMAV) using flight test data. It aims at providing simple FWMAV aerodynamic models that can be used in simulations as well as in nonlinear flight control systems. The undertaken methodology builds on normal aircraft system identification methods and extends these with techniques that are specific to FWMAV model identification. The entire aircraft model identification cycle is discussed covering the set-up and automatic execution of the flight test experiments, the aircraft states, the aerodynamic forces and moments' reconstruction, the aerodynamic model structure selection, the parameter estimation and finally, the model validation. In particular, a motion capturing facility was used to record the flapper's position in time and from there compute the states and aerodynamic forces and moments that acted on it, assuming flap-averaged dynamics and linear aerodynamic model structures. It is shown that the approach leads to aerodynamic models that can predict the aerodynamic forces with high accuracy. Despite less accurate, the predictions of the aerodynamic moments still follow the general trend of the measured moments. Dynamic simulations based on the identified aerodynamic models show flight trajectories that closely match the ground truth spanning a number of flapping cycles. Finally, the dimensional aerodynamic forces and moments' coefficients of two of the identified aerodynamic models are presented.

Online Aerodynamic Model Identification Using a Recursive Sequential Method for Multivariate Splines

Avoiding high computational loads is essential to online aerodynamic model identification algorithms, which are at the heart of any model-based adaptive flight-control system. Multivariate simplex B-spline methods are excellent function approximation tools for modeling the nonlinear aerodynamics of high-performance aircraft. However, the computational efficiency of the multivariate simplex B-spline method must be improved in order to enable real-time onboard applications, for example, in adaptive nonlinear flight-control systems. In this paper, a new recursive sequential identification strategy is proposed for the multivariate simplex B-spline method aimed at increasing its computational efficiency, thereby allowing its use in onboard system identification applications. The main contribution of this new method is a significant reduction of the computational load for large-scale online identification problems as compared to the existing multivariate simplex B-spline methods. The proposed method consists of two sequential steps for each time interval and makes use of a decomposition of the global problem domain into a number of subdomains, called modules. In the first step, the B-coefficients for each module are estimated using a least-squares estimator. In the second step, the local B-coefficients for each module are then smoothened into a single global B-coefficient vector using a linear minimum mean-square errors estimation. The new method is compared to existing batch and recursive multivariate simplex B-spline methods in a numerical experiment in which an aerodynamic model is recursively identified based on data from a NASA F-16 wind-tunnel model.

Towards understanding effects of non-linear flight control system elements on inexperienced pilots

AbstractThis note presents the experimental method and results from a series of desktop simulation tests designed to investigate the manual control characteristics of relatively inexperienced civil pilots; with an average age and experience of 24 years and 66 hours respectively. Increased encroachment into non-linear command gearing was found to make aggressive subjects resort to high levels of crossover regression. The combined effects of rate-limiting and non-linear command gearing were observed only for demanding tasks during which over-control was a typical feature.

Nonlinear generalised dynamic inversion for aircraft manoeuvring control

The nonlinear control problem of aircraft trajectory tracking is tackled in the framework of multiple linear time-varying constrained control using the newly developed paradigm of generalised dynamic inversion. The time differential forms of the multiple constraints encapsulate the control objectives, and are inverted to obtain the reference trajectory-realising control law. The inversion process utilises the Moore–Penrose generalised inverse and the associated nullspace projection, and it predictably involves the problematic generalised inversion singularity. Thus, a singularity avoidance scheme based on a new type of dynamically scaled generalised inverses is introduced that guarantees both asymptotically stable tracking and singularity avoidance. The steady-state closed-loop system allows for two inherently noninterfering control actions working towards a unified goal to exploit the aircraft's control authority over the entire state space. One control action is performed by the particular part of the control law on the range space of the transposed constraint matrix, and it works to impose the prescribed aircraft constrained dynamics. The other control action is performed by the auxiliary part of the control law on the complementary orthogonal nullspace of the constraint matrix, and it provides aircraft's global inner stability using the concept of perturbed feedback linearisation. Numerical simulations of an aggressive multiaxial aircraft coordinated manoeuvre verify the efficacy of designing nonlinear flight control systems via this methodology.

Design and implementation of a robust and nonlinear flight control system for an unmanned helicopter

In this work, we focus on the design and implementation of a robust flight control system for an unmanned helicopter. A comprehensive nonlinear model for an unmanned helicopter system, which is built by our research team at the National University of Singapore, is first presented. A three-layer control architecture is then adopted to construct an automatic flight control system for the aircraft, which includes (1) an inner-loop controller designed using the H ∞ control technique to internally stabilize the aircraft and at the same time yield good robustness properties with respect to external disturbances, (2) a nonlinear outer-loop controller to effectively control the helicopter position and yaw angle in the overall flight envelope, and lastly, (3) a flight-scheduling layer for coordinating flight missions. Design specifications for military rotorcraft set for the US army aviation are utilized throughout the whole process to guarantee a top level performance. The result of actual flight tests shows our design is very successful. The unmanned helicopter system is capable of achieving the desired performance in accordance with the military standard under examination.

Fault tolerant control for near space vehicle: A survey and some new results

A review on fault-tolerant control (FTC) for near space vehicle (NSV) is presented. First, the concept of near space is introduced, the background of NSV is emphasized, and the model characteristics of NSV in faulty case are investigated. Then, a comparison of different existing approaches is briefly carried out, and achievements on the current research in this field are also presented in the view of the practical application. Furthermore, several existing advanced FTC results for nonlinear flight control systems are given. Finally, the recent literature of NSV are pre- sented to provide an overall view of future developments in this area.

Nonlinear Flight Control Systems Design for Hypersonic Vehicles: Results and Open Problems

Air-breathing hypersonic vehicles (HSVs) are regarded as one of the most promising technology for achieving reliable and cost-effective access to space. Notwithstanding the recent successes of NASA's X-43A and AFRL's X-51 experimental vehicles, the design of flight control systems for HSVs is still an open problem, due to the complexity of the dynamics and the unprecedented level of coupling between the airframe and the propulsion system. The slender geometries and light structures required for these aircraft cause significant flexible effects, whereas a strong coupling between propulsive and aerodynamic forces results from the integration of the scramjet engine into the fuselage. Because of the variability of the vehicle characteristics with flight conditions, significant uncertainties affect the vehicle models, which in turn are known to be unstable and non-minimum phase with respect to the regulated output. Finally, the presence of unavoidable constraints on the control inputs and the stringent requirements imposed by hypersonic regimes on the flight envelope render the control design an even harder endeavor, especially concerning controlling the propulsion system. In this paper, we present an account of nonlinear adaptive control techniques that have been developed, in collaboration with the US Air Force Research Laboratories, towards the design of flight control systems for HSVs. In particular, we present several steps leading to the design of a flight control system architecture comprised of a robust adaptive inner-loop controller and a self-optimizing guidance system. Finally, we briefly discuss open problems and current research directions.

Nonlinear Aircraft Loads in Severe Atmospheric Turbulence

The problem of predicting loads due to turbulence for aircraft with nonlinear flight-control systems is reviewed, and related recent research in this area is described. Two alternative approaches to the problem, stochastic simulation and worst-case analysis, are compared and problem areas associated with each are identified. It is concluded that the latter method, which involves a search over a specified family of gust patterns to find the pattern thatcauses maximum load,is appropriate for investigating a limited number of criticalcases, butthat apreliminary simplified method of estimation is called for by computational considerations as a means of identifying such critical cases. A new genetic algorithm is described and proposed as the most reliable method of performing the search. It is illustrated how this search algorithm can be simplified computationally by introducing a weighting factor that penalizesthemorecomplexgustpatterns,astepthatissupportedbymeasurementsoftheamplitudesofgustpatterns in severe turbulence. Nomenclature � A = measure of response found by power spectral density analysis a = empirical constant in Eq. (2) b = empirical constant in Eq. (2) e = energy-reduction factor n = number of component ramp gusts in the discrete gust pattern pi = absolute value of the wavelet coefficient SI = information entropy U = energy of the single-wavelet pattern U� = design gust velocity

Adaptive backstepping-based flight control system using integral filters

A backstepping control design procedure for uncertain nonlinear flight control system expressible in parameter-strict feedback form is presented in this paper. The proposed backstepping procedure, in association with sliding model control technique, exploits the possibility of avoiding, under certain suitable assumptions, the overparameterization problem existing in the classical backstepping process. In particular, a sliding-model-based integral filter is introduced to facilitate the development of the derivation of the virtual inputs, thus reducing the computational load with regard to the standard backstepping procedure. Moreover, in simulations, the control parameters in the resulted controller are optimally tuned using a genetic algorithm so as to show the full potential of the proposed control system.

Nonlinear control concepts for a UA

A reconfigurable flight control method is developed to be implemented on an Unmanned Aircraft (VA), a thirty percent scale model of the Cessna 150. This paper presents the details of the UA platform, system identification, reconfigurable controller design, development, and implementation on the UA to analyze the performance metrics. A Crossbow Inertial Measurement Unit provides the roll, pitch, and yaw accelerations and rates along with the roll and pitch. The 100-400 mini-air data boom from SpaceAge Control provides the airspeed, altitude, angle of attack, and the side slip angles. System identification is accomplished by commanding preprogrammed inputs to the control surfaces and correlating the corresponding variations at the outputs. A Single Network Adaptive Critic, which is a neural network-based optimal controller, is developed as part of a nonlinear flight control system. An online learning neural network is augmented to form an outer loop to reconfigure and supplement the optimal controller to guarantee a "practical stability" for the airplane. This paper also presents some simulations from the hardware-in-the-loop testing and concludes with an analysis of the flight performance metrics for the controller under investigation.