Control Laws For Aircraft Research Articles

Research on longitudinal control technology in the transition section of tilt-rotor aircraft based on linear active disturbance rejection control

To solve the problem of strong coupling and control redundancy in the transition section of the tilt-rotor, this paper presents a longitudinal control method based on linear active disturbance rejection control (LADRC), verified by simulation. Firstly, the nonlinear dynamics model of tilt-rotor aircraft based on the mechanism method is established, and the model’s accuracy is verified by comparing the trim result with the generic tilt-rotor aircraft simulation (GTRS) model verified by flight test in XV-15. Then, the control and redundant rudder surface assignment strategies for tilt-rotor aircraft are proposed. Then, the vertical and vertical control law of tilt-rotor aircraft is designed, the linear expansion state observer (LESO) model is analyzed and derived, and the linear active disturbance rejection controller suitable for attitude loop control of tilt-rotor is designed. The altitude and velocity loops are controlled. Finally, the proposed linear active disturbance rejection control (LADRC) controller is verified by simulation, and the whole process of tilting transition is simulated. The results show that the proposed control strategy and the controller designed in this paper have strong anti-disturbance ability, fast response speed, and high control precision and can be used in the transition section of tilt-rotor aircraft. The vertical direction shall be well controlled.

Integrated design of lateral-directional control law and vertical tail load for civil aircraft

According to new requirements of airworthiness regulation CS25.353, the advanced lateral-directional P-Beta control law design is studied in this paper. The LQR method based on the tracking command model is used to design the lateral-directional control law parameters. At the same time, the vertical tail load model of aircraft for yaw maneuver is introduced into the lateral-directional control law design model. The aim is to build the integrated design of control law and flight load. Finally, the simulation results show that the P-Beta control law can not only improve Dutch roll damping and coordination turn characteristics, but it can also reduce the yaw maneuver load by 18.5%. Thus, the structure weight of the aircraft’s vertical tail can be reduced.

Singular perturbation decomposition-based ride quality control of elastic aircraft

Considering the reduction of ride comfort under wind disturbance, the ride quality control method based on singular perturbation decomposition is proposed. For the dynamic model of elastic aircraft, the singular perturbation theory is used to decouple the model into the rigid-slow subsystem and the flexible-fast subsystem. Considering the additional time-varying disturbance and aerodynamic uncertainty of the rigid subsystem, the disturbance observer is designed to estimate disturbance effect and the neural network is used to deal with model uncertainty. The adaptive robust control is constructed using the composite estimation information as feedforward compensation and the tracking error of pitch rate and normal overload as feedback design. For the flexible subsystem, a nonsingular terminal sliding mode controller is designed to achieve active vibration suppression. The ride quality control law of elastic aircraft is obtained by combining the control inputs of the rigid and flexible subsystems, and the additional normal overload and elastic mode can be quickly restrained and converged. Based on Lyapunov stability analysis, the uniformly ultimate boundedness of the system is proved. The simulation results show that the proposed method can reduce the additional normal overload at the key positions of the aircraft under discrete gust and atmospheric turbulence, and the riding quality of elastic aircraft is effectively improved.

Aeroelasticity Model for Highly Flexible Aircraft Based on the Vortex Lattice Method

With the increasing use of composite materials in aviation, structural aircraft design often becomes limited by stiffness, rather than strength. As a consequence, aeroelastic analysis becomes more important to optimize both aircraft structures and control algorithms. A low computational cost aeroelasticity model based on VLM and rigid-body dynamics is proposed in this work. UAV flight testing is performed to evaluate the accuracy of the proposed model. Two flight sections are chosen to be modeled based on recorded aerodynamic surface control data. The calculated accelerations are compared with recorded flight data. It is found that the proposed model adequately captures the general flight profile, with acceleration peak errors between −6.2% and +8.4%. The average relative error during the entire flight section is 39% to 44%, mainly caused by rebounds during the beginning and end of pull-up maneuvers. The model could provide useful results for the initial phases of aircraft control law design when comparing different control algorithms.

Research on Fault-Tolerant Control of Combined Airframe Damage of Electric Aircraft

General aviation is an important branch of the aviation field. As a green energy aircraft, the electric aircraft is an important component and development direction of general aviation aircraft, and its safety is crucial. In this paper, the aerodynamic and dynamic characteristics of electric aircraft under collision, lightning strikes, and icing conditions are studied, and the dynamic and kinematics models of the aircraft are established by introducing damage factors. The STAR-CCM+ software is used to simulate the aerodynamic force and aerodynamic moment in the case of combined airframe damage. Based on the estimation ability of the L1 adaptive control algorithm for the parameter uncertainty of the controlled object and the automatic adjustment ability of control output, a fault-tolerant control law for electric aircraft is designed in the case of wing damage, horizontal tail damage after a collision, horizontal tail icing, and wing lightning damage. The results show that the control law has good fault-tolerant control ability for combined airframe damage of electric aircraft, and the control system has adaptability, anti-interference, and robustness, which has a good engineering reference significance for flight safety control of other transport aircraft.

Carrier Aircraft Flight Controller Design by Synthesizing Preview and Nonlinear Control Laws

This paper proposes an innovative automatic carrier landing control law for carrier-based aircraft considering complex ship motion and wind environment. Specifically, a strategy is proposed to synthesize preview control with an adaptive nonlinear control scheme. Firstly, incremental nonlinear backstepping control law is adopted in the attitude control loop to enhance the anti-disturbance capability of the aircraft. Secondly, to enhance the glide slope tracking performance under severe sea conditions, the carrier motion is predicted, and the forecasted motion is adopted in an optimal preview control guidance law to compensate influences induced by carrier motion. However, synthesizing the inner-loop and outer-loop control is not that straightforward since the preview control is naturally an optimal control law which requires a state-space model. Therefore, low-order equivalent fitting of the attitude-to-altitude high-order system model needs to be performed; furthermore, a state observer needs to be designed for the low-order equivalent system to supply required states to the landing controller. Finally, to validate the proposed methodology, an unmanned tailless aircraft model is used to perform the automatic landing tasks under variant sea conditions. Results show that the automatic carrier landing system can lead to satisfactory landing precision and success rate even under severe sea conditions.

Incorporating Physical Models for Dynamic Stall Prediction Based on Machine Learning

Unsteady aerodynamic prediction is an important part of modern aircraft safety and control law design. To improve the accuracy and efficiency for unsteady aerodynamic prediction of aircraft at high angles of attack, this paper proposed a machine learning framework based on multifidelity methods. The framework combines the linear dynamic derivative model and the fuzzy neural network model, which can achieve higher prediction accuracy under sparse experimental states. A series of wind-tunnel tests was carried out for the pitching motions of NASA Common Research Model at high angles of attack, to obtain steady and unsteady aerodynamic loads. These experimental data are used to verify the prediction accuracy of the unsteady model in a wide range of oscillation amplitude, frequencies, and mean angles of attack. The results show that the method has good generalization capability for the parameters of interest. At the same time, the comparison with the prediction results only from high-fidelity data shows that the proposed method can effectively reduce the amount of data required for the model of training and improve the modeling robustness to different types of motions.

Co-design of aircraft vertical tail and control laws with distributed electric propulsion and flight envelop constraints

A sequential co-design framework has been developed in a previous study (Nguyen-Van et al., IFAC-PapersOnLine 52(12):514–519, https://doi.org/10.1016/j.ifacol.2019.11.295 , 2019)) to design an aircraft using active differential thrust. Differential thrust is used instead of a rudder to generate the yawing moment. The objective is to dimension in parallel the vertical tail surface area, the electric motor bandwidth and control laws while maintaining imposed handling qualities. This paper focuses on the development of a single step co-design taking into account handling qualities, flight envelop requirements and motor saturation. Additional and compatible optimisation constraints are found based on a sensitivity analysis. It reveals the importance of electric motor bandwidth with respect to aircraft natural stability. The direct co-design leads to an optimised trade-off between vertical tail and electric motor bandwidth.

Research on Air Flight Simulation Control Law of Large Aircraft

The air flight simulator, also known as the variable stability aircraft, can make the dynamic characteristics of the machine change or simulate the response characteristics of different objects in a large range through the variable stability telex system. This paper is devoted to the research of the air flight simulation control law of large aircraft, and designs the five degree of freedom variable stability control law: to control the elevator, rudder and aileron of the basic aircraft, and to design the three-degree-of-freedom variable stability control law. The control of the direct lift surface and the automatic throttle was added, and the five-degree-of-freedom variable stability control law was designed to simulate the trajectory characteristics and velocity characteristics. Finally, the variable stability control law is verified based on the target aircraft simulator.

Design of Composite Feedback and Feedforward Control Law for Aircraft Inertially Stabilized Platforms

The design of the control systems of the inertially stabilized platforms (ISPs) as part of airborne equipment for the majority of aircraft has its peculiarity. The presence of rate gyros in the inertial measurement unit gives the possibility to measure the rotation rate of the ISP base, which is the main disturbance interfering with the ISP accuracy. Inclusion of the feedforward disturbance gain in the control law with the simplest PI feedback significantly improves the accuracy of stabilization by the invariance theory. A combination of feedback and feedforward controllers produces a synergetic effect, thus, improving ISP accuracy. This article deals with the design of the airborne ISP control systems consisting of two stages: the parametric optimization of the PI feedback control based on composite “performance-robustness” criterion and the augmentation of the obtained system with feedforward gain. To prove the efficiency of the proposed control laws, the simulation of the ISP was undertaken. We have used a simulation of the heading-hold system of the commuter aircraft Beaver and the yaw rate output of this closed-loop system we have used as a source of the disturbance. The results of modeling proved the efficiency of the proposed design method.

Extension of the Basic Strapdown INS Algorithms to Solve Polar Navigation Problems

The Earth-centered, Earth-fixed (ECEF) frame does not have any specific points in the near-Earth space; therefore, it can be used as the basic one for navigation in the polar regions. The paper considers a new algorithm for calculation of ECEF coordinates. The proposed algorithm is complementary to the basic strapdown INS algorithm, which functions in its normal mode from aircraft takeoff to landing. In the polar regions, the aircraft control law that supports flights on the point-to-point principle is based on calculation of ECEF coordinates.

Flexible Aircraft Gust Load Alleviation with Incremental Nonlinear Dynamic Inversion

This paper designs an incremental nonlinear dynamic inversion control law for free-flying flexible aircraft, which can regulate rigid-body motions, alleviate gust loads, reduce the wing root bending moment, and suppress elastic modes. By fully exploring the sensor measurements, the model dependency of the proposed control law can be reduced while maintaining desirable robustness, which simplifies the implementation process and reduces the onboard computational load. The elastic states are observed online from accelerometer measurements, with a Padé approximation to model the pure time delay. Theoretical analyses based on the Lyapunov methods and the nonlinear system perturbation theory show that the proposed control has inherent robustness to model uncertainties, external disturbances, and sudden actuator faults. These merits are demonstrated by time-domain simulations in various spatial turbulence and gust fields, as well as by a Monte Carlo study.

Co-design of aircraft vertical tail and control laws using distributed electric propulsion

Abstract Distributed Electric Propulsion is investigated as a way to increase directional control of aircraft and reduce the vertical tail surface area. A co-design approach is presented where a H∞ control method is used to both synthesize longitudinal/lateral control law gains and allocation module while sizing the vertical tail surface area and the propeller actuator bandwidth. The variation of the vertical tail surface area is captured throughout a collection of linearized aircraft systems representing different sizes of the vertical tail reassembled in a Linear Fractional Representation. This approach allows a reduction of 60% of the vertical tail surface area while maintaining desired dynamic behaviors with low actuator bandwidth.

Longitudinal flight control laws for high aspect ratio light utility aircraft

The development of Flight Control Laws (FCLs) for high aspect ratio light utility aircraft have been previously accomplished by considering each control module to be independent. For the longitudinal motion, the FCLs comprise mainly the airspeed control and the flight path angle control. In addition, the altitude control and the vertical speed control have been designed based on the design of flight path angle control. Further investigation is conducted to reveal possibility to design the longitudinal control, i.e. airspeed and flight path angle control by considering the coupling behavior between both controls but still utilizing the same control design strategy that is the advanced Proportional-Integral (PI) control technique. The paper shows the concept and architecture of coupled longitudinal flight control laws, the design strategy chosen, implementation to the aircraft model and investigation of control performance. The results achieved from this research are the architecture and procedure to design longitudinal FCLs. Additionally, knowledge of coupling control behavior between the airspeed and flight path angle via aircraft simulations is acquired.

Stability analysis of aircraft’s sudden damage process with ROA overlapping

PurposeThe paper aims to propose a stability analysis method based on region-of-attraction (ROA) overlapping to evaluate the flight stability of the possible sudden damage process and guarantee flight safety in extreme cases.Design/methodology/approachFirst, according to the two flight conditions before and after damage which the aircraft may encounter, flight dynamical models are built and fitted by polynomial equation for subsequent ROA analysis. And then, the ROA overlapping estimation method based on V-s iteration is presented to complete the stability analysis of such airplane sudden damage process.FindingsFinally, in the presence of control surface damage, case aircraft flight stability is analyzed and simulated to verify the effectiveness of the proposed method.Practical implicationsThe proposed method can be used for stability check during the aircraft control law design, or for further completing the design of the emergency stable controller design.Originality/valueCompared with previous studies on sudden damage process of aircraft, to the best of the authors’ knowledge, this is one of the pioneer studies in which the ROA and ROA overlapping concept has been introduced. This study on stability analysis of aircraft sudden damage process could further develop the theory of emergency stable control.

Distributed Formation Control of Multiple Quadrotor Aircraft Based on Nonsmooth Consensus Algorithms.

The problem of distributed formation control for multiple quadrotor aircraft in the form of leader-follower structure is considered in this paper. Based on a nonsmooth backstepping design, a novel consensus formation control algorithm is proposed and utilized. First, for the position control subsystem, based on the linear quadratic regulator optimal design method, a formation control law for multiple quadrotor aircraft is designed such that the positions of all the quadrotor aircraft converge to the desired formation pattern. The designed formation control law for position systems will generate the desired attitude for the attitude control systems. Second, for the attitude control subsystem described by unit quaternion, by employing the technique of finite-time control and switch control, a global bounded finite-time attitude tracking controller is designed such that the desired attitude can be tracked by the multiple quadrotor aircraft in finite time. Finally, numerical example is performed to demonstrate that all quadrotor aircraft converge to the desired formation pattern in the 3-D-space.

Integrated design of flight control surfaces and laws for new aircraft configurations

Control architecture sizing is a main challenge for new aircraft design like blended wing-body design. This aircraft configuration typically features redundant elevons located at the trailing edge of the wing, acting simultaneously on pitch and roll axes. The problem of integrated design of control surface sizes and flight control laws for an unstable blended wing-body aircraft is addressed here. Latest tools for H∞ non-smooth optimization of structured controllers are used to optimize in a single step the gains for both longitudinal and lateral control laws, and a control allocation module, while minimizing control surfaces total span. Following constraints are ensured: maximal deflection angles and rates for 1) pilot longitudinal pull-up 2) pilot bank angle order and 3) longitudinal turbulence. Using this coupled approach, significant gains in terms of outer elevons span compared to the initial layout are demonstrated, while closed-loop handling qualities constraints are guaranteed.

Initial virtual flight test for a dynamically similar aircraft model with control augmentation system

To satisfy the validation requirements of flight control law for advanced aircraft, a wind tunnel based virtual flight testing has been implemented in a low speed wind tunnel. A 3-degree-of-freedom gimbal, ventrally installed in the model, was used in conjunction with an actively controlled dynamically similar model of aircraft, which was equipped with the inertial measurement unit, attitude and heading reference system, embedded computer and servo-actuators. The model, which could be rotated around its center of gravity freely by the aerodynamic moments, together with the flow field, operator and real time control system made up the closed-loop testing circuit. The model is statically unstable in longitudinal direction, and it can fly stably in wind tunnel with the function of control augmentation of the flight control laws. The experimental results indicate that the model responds well to the operator’s instructions. The response of the model in the tests shows reasonable agreement with the simulation results. The difference of response of angle of attack is less than 0.5°. The effect of stability augmentation and attitude control law was validated in the test, meanwhile the feasibility of virtual flight test technique treated as preliminary evaluation tool for advanced flight vehicle configuration research was also verified.

Параметризация аналитических законов управления боковым движением летательного аппарата

Параметризация аналитических законов управления боковым движением летательного аппарата

A longitudinal flight control law to accommodate sensor loss in the RECONFIGURE benchmark

The feedback gains in state-of-the-art flight control laws for commercial aircraft are scheduled as a function of values such as airspeed, mass, and centre of gravity (CoG). If measurements or estimates of these are lost due to multiple simultaneous sensor failures, the pilot must revert to an alternative control law, or, in the ultimate case, directly command control surface positions. This work develops a robust backup load-factor tracking control law, that does not depend on these parameters, based on application of theory from robust MPC and H2 optimal control. Firstly, the methods are applied with loss only of airdata, and subsequently also with loss of mass and CoG estimates. Local linear analysis indicates satisfactory performance over a wide range of operating points. To keep the aircraft within an acceptable operating region, an outer protection loop is implemented using an override approach, based on ground speed, a model of the trim angle of attack and variation of load factor with respect to angle of attack, and a priori bounds on the wind speed. Finally, the resulting control laws are demonstrated on the nonlinear RECONFIGURE benchmark, which is derived from an Airbus high fidelity, industrially-validated simulator.