Search results

Traditional skid-to-turn (STT) missile control mode is adopted mostly, but with the improvement of requirements for mobility and the emergence of new aerodynamic layout, a bank-to-turn (BTT) control mode gradually shows a greater advantage. However, the BTT missile also has certain defects, for example, when attacking against a maneuvering target and at the last section of guidance, the maximum lifting surface position of the missile needs to be adjusted frequently, thereby increasing the difficulty of control as well as introducing high-frequency noise.

Based on respective characteristics of the two control modes, this paper puts forward a hybrid autopilot design method based on nonlinear dynamic inversion. Firstly, the method converts overload instructions into corresponding angle instructions through the design of hybrid control guidance logic; secondly, based on the nonlinear dynamic inversion algorithm and combined with the fast-changing circuit/slowly changing circuit, a hybrid controller is designed; finally, combined with the missile mathematical model and actuator, it forms a autopilot design closed loop.

The simulation result shows that the non-linear dynamic inverse-based BTT/STT hybrid controller can input a track command well, normal overload and roll angle tracking performance have more advantages than the hybrid controller designed on the basis of classical control method in terms of overshooting and hysteretic characteristics.

The paper puts forward a new BTT/STT hybrid control method which has both the high mobile ability of the BTT missile and the precise control ability of the STT missile, which can adapt to the more complicated fighting environment. And, the method can effectively weaken the impact of the transformation of the control mode on the system.

The purpose of this paper is to find an omnidirectional robust gust response stabilization (GRS) scheme with anti-disturbance and state-limited features.

Disturbance observer and barrier Lyapunov techniques, which can, respectively, estimate the lumped disturbances of the dynamic system in real-time and ensure the middle states within some prescribed ranges according to some flight safety indexes.

In the existing literature, almost all of the GRS controllers are either only for the longitudinal dynamics or only for the latitudinal dynamics. Few studies have considered the gust response alleviation problem with omnidirectional wind disturbance and full aircraft model.

This paper proposes a fresh scheme to deal with a more holistic GRS problem; the disturbance observer based (DOB) barrier Lyapunov backstepping longitudinal controller has been put forward; DOB nonlinear dynamic inversion to handle the multi-input-multi-output lateral dynamics; and to closely connect the two loops of the latitudinal dynamics, a manipulating variable conversion method is proposed.

This paper aims to focus on the sensor fault-tolerant control (FTC) for civil aircraft under exterior disturbance.

First, a three-step cubature Kalman filter (TSCKF) is designed to detect and isolate the sensor fault and to reconstruct the sensor signal. Meanwhile, a nonlinear disturbance observer (NDO) is designed for disturbance estimation. The NDO and the TSCKF are combined together and an NDO-TSCKF is proposed to solve the problem of sensor faults and bounded disturbances simultaneously. Furthermore, an FTC scheme is designed based on the nonlinear dynamic inversion (NDI) and the NDO-TSCKF.

The method is verified by a Cessna 172 aircraft model under bias gyro fault and constant angular rate disturbance. The proposed NDO-TSCKF has the ability of signal reconstruction and disturbance estimation. The proposed FTC scheme is also able to solve the sensor fault and disturbance simultaneously.

NDO-TSCKF is the novel algorithm used in sensor signal reconstruction for aircraft. Then, disturbance observer-based FTC can improve the flight control system performances when the system with faults.

The NDO-TSCKF-based FTC scheme can be used to solve the sensor fault and exterior disturbance in flight control. For example, the bias gyro fault with constant angular rate disturbance of a civil aircraft is studied.

Signal reconstruction for critical sensor faults and disturbance observer-based FTC for civil aircraft are useful in modern civil aircraft design and development.

This is the research paper studies on the signal reconstruction and FTC scheme for civil aircraft. The proposed NDO-TSCKF is better than the current reconstruction filter because the failed sensor signal can be reconstructed under disturbances. This control scheme has a better fault-tolerant capability for sensor faults and bounded disturbances than using regular NDI control.

To design a robust attitude control system for the ascent flight phase of satellite launch vehicles (SLVs).

The autopilot is based on generalized dynamic inversion (GDI). Dynamic constraints are prescribed in the form of differential equations that encapsulate the control objectives, and are generalized inverted using the Moore-Penrose Generalized Inverse (MPGI) based Greville formula to obtain the control law. The MPGI is modified via a dynamic scaling factor for assuring generalized inversion singularity-robust tracking control. An additional sliding mode control (SMC) loop is augmented to robustify the GDI closed-loop system against model uncertainties and external disturbances.

The robust GDI control law allows for two cooperating controllers that act on two orthogonally complement control spaces: one is the particular controller that realizes the dynamic constraints, and the other is the auxiliary controller that is affined in the null control vector, and is used to enforce global closed-loop stability.

Orthogonality of the particular and the auxiliary control subspaces ensures noninterference of the two control actions, and thus, it ensures that both actions work toward a unified goal. The robust control loop increases practicality of the GDI control design.

The first successful implementation of GDI to the SLV control problem.

A low-level controller is critical to the overall performance of multirotor unmanned aerial vehicles. The purpose of this paper is to propose a nonlinear low-level angular velocity controller for multirotor unmanned aerial vehicles in various operating conditions (e.g. different speed and different mode).

To tackle the above challenge, the authors have designed a nonlinear low-level controller taking the actuator dynamics into account. The authors first build the actuator subsystem by combining the actuator dynamics with the angular velocity dynamics model. Then, a recursive low-level controller is developed by designing a high-gain observer to estimate unmeasurable states. Furthermore, a detailed stability analysis is given with the Lyapunov theory.

Simulation tests and real-world flying experiments are provided to validate the proposed approach. In particular, we illustrate the performance of the proposed controller using violent random command test, attitude mode flight and high-speed flight of up to 18.7 m/s in real world. Compared with the classical method used in PX4 autopilot and the estimation-based incremental nonlinear dynamic inversion method, experimental results show that the proposed method can further reduce the control error.

Low-level control of multirotor UAVs is challenging due to the complex dynamic characteristics of UAVs and the diversity of tasks. Although some progress has been made, the performance of existing methods will deteriorate as operating conditions change due to the disregard for the electromechanical characteristics of the actuator.

To solve the low-level angular velocity control problem in various operating conditions of multirotor UAVs, this paper proposes a nonlinear low-level angular velocity controller which takes the actuator dynamics into account.

Flight simulators are one of the noticeable breakthroughs in aerospace engineering. One of the main compartments of flight simulators is its control loading system (CLS). The CLS functions as a generator of virtual aerodynamic control-loads over control columns of a simulator. This paper aims to present the design of a high-fidelity six six degrees of freedom (6DOF) nonlinear CLS for the Boeing-747 aircraft simulator.

An introduction to CLS for flight motion simulators are first recapitulated. Afterward, the commanding devices are explained through schematics available in an engineering sense. This paper then presents in detail, the active control loading strategy and hardware design for the CLS, while also introducing the aerodynamic model structure. The satisfactory computer numerical simulations are presented before the paper ends up in concluding remarks.

The multiple input multiple output (MIMO) 6DOF nonlinear CLS for Boeing-747 flight simulator has been successfully developed. The outcome of computer simulations in real-time verifies practicality of the design strategy. The research presented in this paper could be a simple roadmap for prototyping high-fidelity 6DOF nonlinear CLS for flight motion simulators.

The available control architecture and hardware technologies cannot enable a high-fidelity load realization in a CLS. The existing research has not yet presented a 6DOF nonlinear MIMO CLS architecture along with the underlying controller setup for a high-fidelity load realization. In this paper, the design of a high-fidelity 6DOF nonlinear MIMO CLS for flight simulator of a large transport aircraft has been accomplished.

The purpose of this paper is to develop an attitude control strategy for the reusable boosted vehicle with large angle of attack, and to remove the cross coupling among roll, pitch and yaw channels.

The coordinated gain scheduling control strategy consists mainly of two parts. First, initially ignoring dynamic coupling, single channel gain scheduling controller is designed based on linearized models, respectively. Second, with respect to main channel gain scheduling controller, coordinated scheduling controller is used to generate intentionally cross coupling to partly cancel inter‐channel cross coupling of reusable boosted vehicle.

A coordinated gain scheduling control strategy is presented, and no such analytical solution can be found for the reusable boosted vehicle.

The design idea of coordinated gain scheduling strategy is straightforward in physical concepts and has great value for engineering applications.

Coordinated gain scheduling control strategy is novel in that single channel gain scheduling design does not involve small perturbation linearization and coordinated channel is scheduled.

The purpose of the paper is to design a nonlinear dynamic inversion (NDI) based robust fault-tolerant control (FTC) for aircraft longitudinal dynamics subject to system nonlinearities, aerodynamic parametric variations, external wind disturbances and fault/failure in actuator.

An uncertainty and disturbance estimator (UDE) technique is used to provide estimate of total disturbance enabling its rejection and thereby achieving robustness to the proposed NDI controller. As needed in the NDI design, the successive derivatives of the output are obtained through an UDE robustified observer making the design implementable. Further, a control allocation scheme consigns control command from primary actuator to the secondary one in the event of fault/failure in the primary actuator.

The robustness is achieved against the perturbations mentioned above in the presence of actuator fault/failure.

Lyapunov analysis proves practical stability of the controller–observer structure. The efficacy and superiority of the proposed design has been demonstrated through Monte-Carlo simulation.

Unlike in many FTC designs, robustness is provided against system nonlinearities, aerodynamic parametric variations, external wind disturbances and sinusoidal input disturbance using a single control law which caters for fault-free, as well as faulty actuator scenario.

The purpose of this paper is to analyze and compare the performance of several different UAV trajectory tracking algorithms in normal and abnormal flight conditions to investigate the fault‐tolerant capabilities of a novel immunity‐based adaptive mechanism.

The evaluation of these algorithms is performed using the West Virginia University (WVU) UAV simulation environment. Three types of fixed‐parameter algorithms are considered as well as their adaptive versions obtained by adding an immunity‐based mechanism. The types of control laws investigated are: position proportional, integral, and derivative control, outer‐loop nonlinear dynamic inversion (NLDI), and extended NLDI. Actuator failures on the three channels and increased turbulence conditions are considered for several different flight paths. Specific and global performance metrics are defined based on trajectory tracking errors and control surface activity.

The performance of all of the adaptive controllers proves to be better than their fixed parameter counterparts during the presence of a failure in all cases considered.

The immunity inspired adaptation mechanism has promising potential to enhance the fault‐tolerant capabilities of autonomous flight control algorithms and the extension of its use at all levels within the control laws considered and in conjunction with other control architectures is worth investigating.

The WVU UAV simulation environment has been proved to be a valuable tool for autonomous flight algorithm development, testing, and evaluation in normal and abnormal flight conditions.

A novel adaptation mechanism is investigated for UAV control algorithms with fault‐tolerant capabilities. The issue of fault tolerance of UAV control laws has only been addressed in a limited manner in the literature, although it becomes critical in the context of imminent integration of UAVs within the commercial airspace.

The air‐breathing hypersonic vehicle (AHV) includes intricate inherent coupling between the propulsion system and the airframe dynamics, which results in an intractable nonlinear system for the controller design. The purpose of this paper is to propose an H∞ control method for AHV based on the online simultaneous policy update algorithm (SPUA).

Initially, the H∞ state feedback control problem of the AHV is converted to the problem of solving the Hamilton‐Jacobi‐Isaacs (HJI) equation, which is notoriously difficult to solve both numerically and analytically. To overcome this difficulty, the online SPUA is introduced to solve the HJI equation without requiring the accurate knowledge of the internal system dynamics. Subsequently, the online SPUA is implemented on the basis of an actor‐critic structure, in which neural network (NN) is employed for approximating the cost function and a least‐square method is used to calculate the NN weight parameters.

Simulation study on the AHV demonstrates the effectiveness of the proposed H∞ control method.

The paper presents an interesting method for the H∞ state feedback control design problem of the AHV based on online SPUA.