The control of the hypersonic glide vehicle (HGV) during re-entry is confronted with the input saturation and unknown dynamical uncertainty problem. A nonlinear anti-windup (AW) compensation control method is designed to deal with the nonlinear control problem caused by the saturation of the control surfaces of the HGV. The scheme transforms the problem of saturation control into the optimization problem through nonlinear matrix inequality (NMIL) constraints, which is solved by particle swarm optimization (PSO). The saturation situation can be relieved within a short time and the satisfactory control performance is attained by the method. Then, a novel sigmoid function tracking differentiator based disturbance observer (STDDO) is proposed to estimate parameter uncertainties and external disturbances for the problem of dynamical uncertainties. This method does not need the priori information about the bounds of disturbances and has global fast convergence property. Finally, the simulation results show the effectiveness and robustness of the proposed control scheme.

A nonlinear anti-windup(AW) control compensator based on improved cooperative optimization algorithm(ICOA) is proposed to deal with the control problem caused by the saturation of the control surfaces of the hypersonic vehicle(HSV) during re-entry. Firstly, saturated link is added in the system controller, and the model of AW compensator is designed according to the HSV system equations. Then, the nonlinear L2 gain structure is introduced with the purpose of optimizing the performance of system, which transforms the problem of saturation control into the optimization problem through nonlinear matrix inequality (NMIL) constraints. The optimization problem is solved by an ICOA. Finally, the simulation results show the effectiveness of the proposed method. The saturation situation can be relieved within a short time and the satisfactory control performance is attained by the presented method.

In aircraft flight test phase, in order to verify whether the aircraft has reached the reliability requirements, we need to assess its reliability. Firstly, in this paper we review the reliability estimation method under multi-profiles that can be used to evaluate the reliability of aircraft during the flight test process and analyze its shortcoming. Then, in order to expand the amount of data, we describe how to incorporate the environment conversion factors into the reliability estimation method under multi-profiles. By the way, taking into account the varied population-environment feature of data, an integrated evaluation model of flight test of reliability was established. This model is making up the weak points of previous models, and improving the accuracy of reliability evaluation. Finally, this method applied in a real engineering example is described.

BIT's false alarm is one of the most important reasons of restricting testability's development. In order to decrease BIT's false alarm and develop BITE's performance, this paper discusses present situation of BIT's false alarm at first, then introduces several advanced models of reducing false alarm which are neural network, support vector machine, fuzzy clustering analysis and hidden Markov model. After then, this paper compares the advantages and disadvantages of these models about reducing BIT's false alarm. Synthesizing the advantages of SVM, FCA and HMM, this paper gives an intelligent model based on SVM_FCA_HMM to decrease false alarm, and analyzes theoretically this model's functionary and efficiency. Finally, this paper applies the method in the aero-engine's digital electronic control system and discusses its rationality and applicability, and proves that this model can be applied generally.

This paper presents a method for the construction of reliability, maintainability and supportability (RMS) parameters system for support system. In the beginning, support system is defined from consists and function, support mission profile and system feature analysis. Then with product class and the selection criteria of RMS parameters, a complete RMS parameters system will be built with the selection of reliability parameters suitable for the system from different aspects. Furthermore, according to the need of support system, combined with the characteristic of the study object, one or more integrated parameters related with RMS can be acquired. In the end, this paper proves its operability combined with the process of building RMS parameters system of one launch site system.