Nonlinear Aircraft Model Research Articles

A Control-Theoretic Study on Iterative Solution to Control Allocation for Over-Actuated Aircraft

The approach of feedback iteration in control theory is considered for solving the problem of control allocation algorithm for the case of three moments in this paper. The core of our novel idea is to introduce a closed-loop feedback iteration for obtaining an available solution to control allocation, which is referred to as control theoretic control allocation (CTCA). Two types of CTCA are investigated in this paper: linear case and nonlinear case, where, in the former, the control Lyapunov function (Lyapunov optimizing) and conjugate gradient are introduced, and in the latter, a nonlinear feedback theory is utilized. Thus, the control Lyapunov function and Lyapunov optimizing control allocation (CLF/LOCA), conjugate gradient control allocation (CGCA), and nonlinear CTCA (NCTCA) are proposed in this paper. The performance of each algorithm is evaluated for computational efficiency and accuracy by use of linear or nonlinear aircraft models, respectively. The simulation results show that the proposed CTCA algorithm will terminate in finite steps.

Required Moment Sets: Enhanced Controllability Analysis for Nonlinear Aircraft Models

Attainable moment sets (AMS) are a powerful method to assess aircraft controllability. However, as attainable moment sets only take into account the achievable moment set of the control effectors, they do not assess the required moment set to fulfill the aircraft mission requirements. This paper proposes to calculate a corresponding required moment set (RMS) which defines a set of moments sufficient for fulfilling aircraft controllability requirements in the mission flight envelope. The paper applies the required moment set approach to a nonlinear simulation model of an electric vertical take off vehicle (eVTOL) transition drone in hover. By comparing the required moment set to the AMS of the aircraft model, moment set margins are derived and used to assess the controllability of the considered aircraft. The results indicate that the combined evaluation directly identifies critical moment channels and margins, which is advantageous when compared to a pure AMS-based evaluation. The proposed approach enables the execution of simulation-based assessments in aircraft design and flight control development. In the early stages of aircraft design, required moment sets can support sizing, positioning and tilting of control effectors (e.g., propulsive elements) to fit the AMS to the actual required force and moment set for the specific system.

Adaptive fuzzy direct lift control of aircraft carrier-based landing

Considering the precise landing demand for carrier-based aircraft flight control, this paper proposes an adaptive fuzzy landing control method for the strong time-varying, parameter uncertainty and various complex coupling in the actual state aircraft model. This method is applied to the flap channel to achieve direct lift control, and the fuzzy system is utilized to approximate the six-degree-of-freedom nonlinear system model of a carrier aircraft that is difficult to accurately describe, to achieve accurate tracking of the landing glideslope, and improve the landing accuracy. Lyapunov method was used to judge the stability of the adaptive fuzzy control algorithm. During the simulation, the airwake and deck motion disturbance were introduced to simulate the landing environment of the aircraft. The effectiveness of the landing control system was verified by the Matlab software. Monte-Carlo random test was utilized to carry out the landing point accuracy for the conventional control scheme and the adaptive fuzzy direct lift control scheme respectively. Through the response curve and landing point statistical results, it is confirmed that the direct lift control scheme has better control effect, and the landing precision has improvement compared with the conventional control scheme.

Back-stepping Active Disturbance Rejection Control for Attitude Control of Aircraft Systems Based on Extended State Observer

Robust flight control laws based on back-stepping technology and ADRC method are designed for attitude control of a non-linear aircraft system. First, non-linear aircraft model is introduced and converted to standard equation of state. Then, the extended state observer is applied to estimate the unknown variables, the homologous ADRC is designed to ensure the state variables of the CLS to astringe to the reference state. Next, the stability of ESO and ADRC are analyzed and proven theoretically. At last, the effectiveness of this method is illustrated by extensive comparative simulations. The results acquired from simulation attest that ADRC can achieve better control performance than PID and SMC method.

An improved online system identification method for tiltrotor aircraft

An online system identification method for tiltrotor aircraft flight dynamics modeling is developed in this paper. A general parametric flight dynamics model which can be used for online identification purpose in the three flight modes of tiltrotor aircraft (i.e., helicopter mode, conversion mode, and fixed-wing mode) is developed, and an unideal noise model is also introduced in order to minimize the parameter identification error caused by measurement noise. An adaptive model structure identification algorithm is established by introducing local accuracy criteria, so that a time variant optimal model structure is obtained in real time. A weighted recursive least squares algorithm is established for parameter identification, and a time variant weighting matrix is designed to provide best numerical performance in the identification procedure at any time. A comprehensive nonlinear flight dynamics model of a notional tiltrotor aircraft is used to generate simulation flight test data in all flight modes, and the identification results of the notional tiltrotor aircraft in fixed wing mode, helicopter model as well as conversion mode are obtained and verified. Finally, the flight dynamics model of an unmanned tiltrotor aircraft in helicopter mode is identified by using the real flight test data. The identification results show that the online system identification method for tiltrotor aircraft developed in this paper is effective, efficient and robust.

DYNAMIC MODELING AND STABILITY CONTROL OF FOLDING WING AIRCRAFT

During the deformation process, the dynamic modeling of the folding-wing aircraft presents the characteristics of multi-rigid、multi-degree of freedom and strong nonlinearity. At the same time, parameters such as aerodynamics/torque, pressure center, centroid and moment of inertia will also change greatly, which will seriously affect Flight stability. Therefore, this paper will mainly study the multi-rigid dynamics modeling and deformation stability control of the folding wing aircraft. The multi-rigid dynamic model of the folding wing aircraft is established based on the Kane method with the additional force and moment expressions. The functional relationship between the aerodynamic parameters and the folding angle is fitted through aerodynamic calculations. and the longitudinal dynamic characteristics of the aircraft at different folding angular speeds are analyzed. It is shown that the speed, height and pitch angle of the folding wing aircraft will change during the deformation process by analyzing the longitudinal dynamic characteristics, and the aircraft cannot maintain stable flight. A stability control method is proposed for the deformation process of the folding-wing aircraft based on the active disturbance rejection control theory. The nonlinear terms, coupling terms and parameter time-varying terms are regarded as the total internal and external disturbances in the longitudinal nonlinear dynamic model of the folding-wing aircraft, using the extended state observer to estimate and compensate the total disturbance in real time. The PD controller is proposed for compensated systems to realize decoupling control of speed channel and height channel. The stability of the system is proved by Lyapunov stability theory, and mathematical simulation is used to verified the stability of the folding wing aircraft. The simulation results show that the stability controller based on the active disturbance rejection control theory can solve the problems of strong nonlinearity and time-varying parameters caused by aircraft deformation, and ensure the high-precision and stable control of the aircraft.

Robust output feedback control of fixed-wing aircraft

PurposeThis paper aims to devise a robust controller for the non-linear aircraft model using output feedback control topology in the presence of uncertain aerodynamic parameters.Design/methodology/approachFeedback linearization-based state feedback (SFB) controller is considered along with a robust outer loop control which is designed using Lyapunov’s second method. A high-gain observer (HGO) in accordance with the separation principle is used to implement the output feedback (OFB) control scheme. The robustness of the controller and observer is assessed by introducing uncertain aerodynamics coefficients in the dynamic model. The proposed scheme is validated using MATLAB/SIMULINK.FindingsThe efficacy of the proposed scheme is authenticated with the simulation results which show that HGO-based OFB control achieves the SFB control performance for a small value of the high-gain parameter in the presence of uncertain aerodynamic parameters.Originality/valueA HGO for the non-linear model of aircraft with uncertain parameters is a novel contribution which could be further used for the unmanned aerial vehicles autopilot, flight trajectory tracking and path following.

Identification and Modeling of the Airbrake of an Experimental Unmanned Aircraft

This paper presents the modeling, system identification, simulation and flight testing of the airbrake of an unmanned experimental aircraft in frame of the FLEXOP H2020 EU project. As the aircraft is equipped with a jet engine with slow response an airbrake is required to increase deceleration after speeding up the aircraft for flutter testing in order to remain inside the limited airspace granted by authorities for flight testing. The airbrake consists of a servo motor, an opening mechanism and the airbrake control surface itself. After briefly introducing the demonstrator aircraft, the airbrake design and the experimental test benches the article gives in depth description of the modeling and system identification referencing also previous work. System identification consists of the determination of the highly nonlinear (saturated and load dependent) servo actuator dynamics and the nonlinear aerodynamic and mechanical characteristics including stiffness and inertia effects. New contributions relative to the previous work are a unified servo angular velocity limit model considering opening against the load or closing with it, the detailed construction and evaluation of airbrake normal and drag force models considering the whole deflection and aircraft airspeed range, the presentation of a unified aerodynamic - mechanic nonlinearity model giving direct relation between airbrake angle, dynamic pressure and servo torque and the transfer function-based modeling of stiffness and inertial effects in the mechanism. The identified servo dynamical model includes system delay, inner saturation, the aforementioned load dependent angular velocity limit model and a transfer function model. The servo model was verified based-on test bench measurements considering the whole opening angle and dynamic load range of the airbrake. New, unpublished measurements with gradually increasing servo load as the servo moves are also considered to verify the model in more realistic circumstances. Then the full airbrake model is constructed and tested in simulation to check realistic behavior. In the next step the airbrake model integrated into the nonlinear simulation model of the FLEXOP aircraft is tested by flying simulated test trajectories with the baseline controller of the aircraft in software-in-the-loop (SIL) Matlab simulation. First, the standalone airbrake simulation is compared to the SIL results to verify flawless integration of airbrake model into the nonlinear aircraft simulation. Then deceleration times with and without airbrake are compared underlining the usefulness of the airbrake in the test mission. Finally, real flight data is used to verify and update the airbrake model and show the effectiveness of the airbrake.

A High Precision Adaptive Back-Stepping Control Method for Morphing Aircraft Based on RBFNN Method

To overcome the uncertainties of the nonlinear model of a morphing aircraft, this paper presents a high-precision adaptive back-stepping control method based on the radial basis function neural network (RBFNN). Firstly, based on the analysis of static and dynamic aerodynamic parameters of the morphing aircraft, its nonlinear control law is designed by using the conventional back-stepping method. The RBFNN is introduced to approximate online the uncertain terms of the nonlinear control law so as to improve its robustness. The robust term is designed to eliminate the approximation error caused by the RBFNN. Secondly, the tracking differentiator is designed through solving the virtual control variables, thus solving the "differential expansion" problem existing in the traditional back-stepping method. The Lyapunov stability analysis proves that our method can ensure that the tracking error of a closed-loop system converges finally and that its signals are uniformly bounded. Finally, the digital simulation model of the morphing aircraft is established with the MATLAB/Simulink; our method is compared with the conventional back-stepping control method. The simulation results show that our method has a higher control precision and stronger robustness.

Black-box Modeling for Aircraft Maneuver Control with Bayesian Optimization

This paper proposes a new method of designing a data-driven controller for aircraft maneuver. Assuming that we do not have knowledge of the controller and the controlled aircraft, we propose a controller design with explorations of the control inputs and their responses from the aircraft. Specifically, we utilize Bayesian optimization (BO) with Gaussian process (GP) regression for black-box modeling of the aircraft responses from the explored controls, which are selected as samples to experiment with BO. We tested the proposed controller with a rigid six degrees of freedom (6DoF) nonlinear aircraft model by varying the kernel structures of the GP regressions. Our proposed method shows shorter flight times and smaller deviations navigating fixed waypoints compared to the tuned Proportional Integral Derivatives (PID) controller. The proposed controller can be an alternative to PID control, particularly when both controller structure and controlled plant model information are unknown.

Integrated optimization of control surface layout for gust load alleviation

Considering active gust load alleviation (GLA) during aircraft design offers great potential for structural weight savings. The effectiveness of a GLA control system strongly depends on the layout of available control surfaces, which is investigated in this article. For the purpose of wing load reduction, a concurrent optimization of controller gains and aileron geometry parameters is carried out. To that end, an efficient update routine for the nonlinear model of a large-scale flexible aircraft with unsteady aerodynamics is presented. Compared to a GLA system using the original aileron configuration, a 9% performance improvement is achieved. Furthermore, a trade-off study is carried out which enables a target-oriented balancing between individual load channels. The significant influence of aileron size and position on overall GLA performance is demonstrated and hence a consideration during the preliminary aircraft design process is recommended.

An investigation of a high-speed ground manoeuvre under optimal control

This paper studies the behaviour of a nonlinear aircraft model under optimal control for aircraft ground manoeuvres, specifically for high-speed runway exits. The aircraft's behaviour on the ground is captured by a fully parameterised 6-DOF nonlinear model, which is developed in this work to model the effects of braking through a combined slip tyre model. A pre-defined cost function is minimised using a generalised optimal control algorithm to obtain an optimal control sequence for a particular manoeuvre-cost function combination. In this paper, three scenarios are investigated for a 45-degree high-speed runway exit: the first control sequence minimises the distance between the aircraft's centre of gravity and the runway centreline; the second maximises the distance travelled by the aircraft during the 20 s of simulation time; the third minimises tyre wear. For each scenario, the generalised optimal control algorithm provides the best possible control inputs. The dynamic response of the aircraft throughout the turn is shown to be dominated by its inertia, which suggests that future controllers will need to begin executing a turn far in advance of entering the corner. The results also provide a benchmark against which the effectiveness of future real-time controllers may be judged.

Impact of spanwise non-uniform discrete gusts on civil aircraft loads

ABSTRACTThe drive for increasing flight efficiency is resulting in wing designs that are of higher aspect ratios, lower in weight, increased wingspans and, consequently, require greater attention in the disciplines of aeroelastics and loads. This trend in aircraft design, along with past research experience with flexible aircraft, motivate a review of assumptions in gust models; especially, that of the gust maintaining a uniform spanwise profile. In this paper, the authors investigate the use of spanwise varying 1−cos gust models for loads prediction using a non-linear aeroelastic model of a conventional large transport aircraft. The comparison between a test case using conventional uni-dimensional approach and another, using multidimensional gusts, illustrates the impact of stepping away from traditional discrete tuned gust processes and adding a spanwise varying gust component. A methodology for processing and analysing the loads data arising due to the added dimension is also developed and both envelope and correlated loads are considered. Gust characteristics and resulting load factor are, respectively, considered for comparison between the two models, as both metrics define realistic gust encounters. In this case, it has been shown that spanwise variation of gust profiles leads to lower envelope loads if viewed in terms of conventional gust gradients. However, higher envelope loads are found if the maximum load factors are matched.

Aircraft pitch control tracking with sliding mode control

Sliding mode control (SMC) is one of the robust and nonlinear control methods. An aircraft flying at high angles of attack is considered nonlinear due to flow separations, which cause aerodynamic characteristics in the region to be nonlinear. This paper presents the comparative assessment for the flight control based on linear SMC and integral SMC implemented on the nonlinear longitudinal model of a fighter aircraft. The controller objective is to track the pitch angle and the pitch rate throughout the high angles of attack envelope. Numerical treatments are carried out on selected conditions and the controller performances are studied based on their transient responses. Obtained results show that both SMCs are applicable for high angles of attack.

Development of Nonlinear Flight Mechanical Model of High Aspect Ratio Light Utility Aircraft

The implementation of Flight Control Law (FCL) for Aircraft Electronic Flight Control System (EFCS) aims to reduce pilot workload, while can also enhance the control performance during missions that require long endurance flight and high accuracy maneuver. In the development of FCL, a quantitative representation of the aircraft dynamics is needed for describing the aircraft dynamics characteristic and for becoming the basis of the FCL design. Hence, a 6 Degree of Freedom nonlinear model of a light utility aircraft dynamics, also called the nonlinear Flight Mechanical Model (FMM), is constructed. This paper shows the construction of FMM from mathematical formulation, the architecture design of FMM, the trimming process and simulations. The verification of FMM is done by analysis of aircraft behaviour in selected trimmed conditions.

Time-varying linear control for tiltrotor aircraft

Tiltrotor aircraft have three flight modes: helicopter mode, airplane mode, and transition mode. A tiltrotor has characteristics of highly nonlinear, time-varying flight dynamics and inertial/control couplings in its transition mode. It can transit from the helicopter mode to the airplane mode by tilting its nacelles, and an effective controller is crucial to accomplish tilting transition missions. Longitudinal dynamic characteristics of the tiltrotor are described by a nonlinear Lagrange-form model, which takes into account inertial/control couplings and aerodynamic interferences. Reference commands for airspeed velocity and attitude in the transition mode are calculated dynamically by visiting a command library which is founded in advance by analyzing the flight envelope of the tiltrotor. A Time-Varying Linear (TVL) model is obtained using a Taylor-expansion based online linearization technique from the nonlinear model. Subsequently, based on an optimal control concept, an online optimization based control method with input constraints considered is proposed. To validate the proposed control method, three typical tilting transition missions are simulated using the nonlinear model of XV-15 tiltrotor aircraft. Simulation results show that the controller can be used to control the tiltrotor throughout its operating envelop which includes a transition flight, and can also deal with vertical gust disturbances.

Fault-Tolerant Flight Control Based on Two-Stage Approach and Interior Point Method

In the field of flight control, it is an important issue to assure the flight performance even if some actuators fail. In the viewpoint of mathematics, aircraft dynamics can be described as an input affine system, so a two-stage control approach for the affine system is adopted here. At the first stage, to generate virtual inputs related to the intended maneuver, nonlinear inverse control technique is adopted. At the second stage, an actuator reassignment problem how to achieve necessary virtual inputs by remaining available actuators is considered. Taking into account the physical limits of the remaining actuators as well as those of the aircraft structure, this problem is organized as a quadratic programming problem solved on-line to generate reference signals for remaining available actuators. The proposed solution method for this formulated programming problem is based on an interior point method. Numerical simulation results based on a nonlinear aircraft model show that the proposed approach can provide an acceptable flight performance with some actuators failure.

Design and Hardware Implementation of Autopilot Control Laws for MAV Using Super Twisting Control

In this paper we present the design and implementation of autopilot tracking control law for Micro Aerial Vehicle using the second order sliding mode approach. The inner loop attitude tracking control design is carried out using output feedback based second order sliding mode technique, to ensure finite time convergence of the tracking error dynamics. While addressing tracking control of a time varying reference signal, it is important to investigate the stability characteristics of the internal dynamics to ensure perfect tracking. This paper mainly addresses the output tracking control problem for a MAV and investigate the stability characteristics of the longitudinal zero dynamics during tracking. We have proposed a stability proof based on Lyapunov theory to analyze the stability of the MAV longitudinal zero dynamics during tracking. A nonlinear aircraft model obtained using aerodynamic derivatives of The Blackkite 300 mm wingspan fixed MAV is used for both control design and as well as to verify its performance against the classical control methods. Extensive hardware in-loop simulation results of the proposed control algorithm implemented on the commercially available PX4 based Pixhawk autopilot board are also presented here.

On the internal stability of non‐linear dynamic inversion: application to flight control

Aircraft are highly non-linear systems, but flight control laws are traditionally designed from a set of linearised models. Due to the application of linear control laws on a non-linear system, the real performance ability of the aircraft is not fully utilised. In addition, in adverse situations like near stall, the aircraft develops significant non-linearities, and linear control laws do not perform well. This study therefore considers the design of a longitudinal flight controller for a fixed-wing aircraft using non-linear dynamic inversion technique or, in terms of control theory, partial exact feedback linearisation. A novel contribution of this study is the proposed combination of three different automatic flight controllers that provide complete 3-DOF longitudinal control. A detailed analysis of the internal dynamics for each controller is also presented. It has been shown that for each controller the internal dynamics are stable. This makes the controller suitable for various flight conditions. The aims of these flight controllers are threefold. First, to provide control of the flight path angle by tracking the pitch angle and the angle of attack. Second, to provide high attitude (pitch up or down) manoeuvrability. Finally, to provide automatic adverse attitude recovery of the aircraft in situations like stall, the switching strategy between the controllers are also discussed. A simulation carried out on a non-linear model of a multi-role fighter aircraft verified the proposed theoretical results confirming the suitability of the controllers.

Type-2 fuzzy adaptive sliding mode control of hypersonic flight

This paper proposes a type-2 fuzzy adaptive sliding mode control scheme for tracking control of hypersonic aircraft with uncertainties. This method uses full-state feedback to linearize the nonlinear model of hypersonic aircraft. Combining the interval type-2 fuzzy approach and adaptive sliding mode control keeps the system stable in the existence of uncertain parameters. For rapid stabilization of the system, the adaptive laws are designed using a direct constructive Lyapunov analysis together with a well-established type-2 fuzzy logic control. Simulation test results indicate that the proposed control scheme provides enhancement of robustness to parametric uncertainty and improvement in tracking performance of the hypersonic aircraft.