In this paper, a model reference adaptive control (MRAC) augmentation method of a linear controller is proposed for air-breathing hypersonic vehicle (AHV) during inlet unstart. With the development of hypersonic flight technology, hypersonic vehicles have been gradually moving to the stage of weaponization. During the maneuvers, changes of attitude, Mach number and the back pressure can cause the inlet unstart phenomenon of scramjet. Inlet unstart causes significant changes in the aerodynamics of AHV, which may lead to deterioration of the tracking performance or instability of the control system. Therefore, we firstly establish the model of hypersonic vehicle considering inlet unstart, in which the changes of aerodynamics caused by inlet unstart is described as nonlinear uncertainty. Then, an MRAC augmentation method of a linear controller is proposed and the radial basis function (RBF) neural network is used to schedule the adaptive parameters of MRAC. Furthermore, the Lyapunov function is constructed to prove the stability of the proposed method. Finally, numerical simulations show that compared with the linear control method, the proposed method can stabilize the attitude of the hypersonic vehicle more quickly after the inlet unstart, which provides favorable conditions for inlet restart, thus verifying the effectiveness of the augmentation method proposed in the paper.

A distributed encirclement points deployment scheme for a group of autonomous mobile agents is addressed in this paper. Herein, each agent can measure its own azimuth related to the common target and can at least communicate with its two adjacent neighbors. Given its space-cooperative character, the encirclement points deployment problem is formulated as the coverage control problem on a circle. The measurement range of azimuth sensor is taken into consideration when doing problem formulation, which is closer to the facts in real-world applications. Then, the fully distributed control protocols are put forward based on geometric principle and the convergence is proved strictly with algebraic method. The proposed control protocols can steer the mobile agents to distribute evenly on the circle such that the coverage cost function is minimized, and meanwhile the mobile agents’ spatial order on the circle is preserved throughout the systems’ evolution. A noteworthy feature of the proposed control protocols is that only the azimuths of a mobile agent and its two adjacent neighbors are needed to calculate the mobile agent’s control input, so that the control protocols can be easily implemented in general. Moreover, an adjustable feedback gain is introduced, and it can be employed to improve the convergence rate effectively. Finally, numerical simulations are carried out to verify the effectiveness of the proposed distributed control protocols.

Aimed at the poor performance of guidance algorithms designed based on a linearized model in active defense under large leading angle deviation, both-way and one-way cooperative sliding mode guidance algorithms based on the smooth nonsingular terminal sliding mode for the defense missile are proposed. The relative kinematics and linearized models of the target, the active defense missile, and the attacking missile are established. In the design process, two smooth nonsingular terminal sliding mode surfaces are constructed based on zero-effort miss distance and zero-effort velocity, as well as their integral values. A tracking differentiator is introduced for excessive initial command deviation to meet the overload constraints of the active missile and the target. The sensitivity of guidance law to the estimated time-to-go error is reduced, and the target is allowed to perform an independent evasive maneuver. The effectiveness of the proposed guidance strategy is verified by numerical simulation and compared to the existing guidance strategies, the high accuracy, anti-chattering, and strong robustness of the proposed guidance algorithm are verified.

For hypersonic-vehicle fairing separation, the reliable separation of the fairing and rapid suppression of the separation disturbance are required in scenarios where there is high dynamic pressure and there are abrupt changes in the aerodynamic shape of the aircraft as well as overall parameters. A switching-based control method is proposed in this paper for the interference suppression of hypersonic-vehicle fairing separation, aiming to ensure the stability of the control system in cases of abrupt changes in the controlled object. First, an unsteady-flow calculation method is adopted to clarify the aerodynamic interference characteristics of the aircraft in fairing separation, and the aerodynamic calculation model of the hypersonic-vehicle fairing-separation process is established; then, two states of the aircraft, with and without fairing, are transformed into a subsystem with switching characteristics. The attitude-stability problem in the fairing-separation process is transformed into a problem of solving the arbitrary switching of the switched system by using the linear matrix inequality (LMI) approach. Meanwhile, the control system for interference suppression of the fairing separation is designed by considering the aerodynamic interaction of the fairing movement with the aircraft during the separation. The simulation results showed that the designed control system can realize the reliable separation of the fairing under high dynamic pressure and effectively suppress interference in the separation process.

Hypersonic cruise missile always adopts the configuration of waverider body with the restraint of scramjet. As a result, the lateral motion exhibits serious coupling, and the controller design of the lateral lateral system cannot be conducted separately for yaw channel and roll channel. A multiple input and multiple output optimal control method with integrators is presented to design the lateral combined control system for hypersonic cruise missile. A hypersonic cruise missile lateral model is linearized as a multiple input and multiple output plant, which is coupled by kinematics and fin deflection between yaw and roll. In lateral combined controller, the integrators are augmented, respectively, into the loop of roll angle and lateral overload to ensure that the commands are tracked with zero steady-state error. Through simulation, the proposed controller demonstrates good performance in tracking the command of roll angle and lateral overload.

Covariance analysis describing function technique is a conventional method to solve the performance analysis of the nonlinear missile guidance system. Aiming at the faultiness of covariance analysis describing function technique and its improved method, this paper proposes a new method to solve the mean and covariance transfer issue by using the high-degree spherical-radial cubature rule. In terms of simulation models, the nonlinear missile guidance system contains the saturation link to limit the lateral acceleration command and the initial heading error is introduced. In addition, the zero-mean target evasive maneuver with exponentially-decaying autocorrelation function is chosen as the input of the missile guidance system. Furthermore, compared with the existing methods, the proposed method has higher computational efficiency, accuracy and better applicability in high-dimensional nonlinear systems by theoretical derivation and simulation results from the mean and covariance values of relative lateral separation. Finally, the presented analysis method can be used as a general method for the design and statistical analysis of multidimensional nonlinear systems.