Forming and keeping fast fly-around under constant thrust

A new switching control algorithm under constant thrust is designed for the chaser fast flying around the target spacecraft along a specified fly-around trajectory. The switching control laws are obtained based on the acceleration sequences and the on time of thrusters which can be computed by the time series analysis method. The perturbations and fuel consumptions are addressed during the computation of the on time of thrusters. Furthermore, the relative position parameters of the target spacecraft are obtained by using the vision measurement and the target fly-around positions are calculated through the isochronous interpolation method. The change of the relative position and the relative velocity of the chaser during the constant thrust fast fly-around are presented through simulation example. It is proved that, with the switching control laws, the chaser will fast fly around the target spacecraft along the specified fly-around trajectory.     

Measurements of heat transfer and pressure in a trailing edge cavity of a turbine blade

 An experimental investigation is conducted to obtain the heat transfer and pressure drop data for an integral trailing edge cavity test section that simulates a novel turbine blade's internal cooling passage with bleed holes. Local heat transfer is measured on both the suction and pressure sides by a transient liquid crystal technique, while pressures at six positions are recorded by pressure calibrators. Moreover, flow characteristic and its effect on heat transfer are analyzed for conditions with or without bleed flow. The experimental results show that, in the cases with bleed flow, local heat transfer on the pressure side exceeds that on the suction side in the first and second channels. In the cases without bleed flow, in the first and third channels, local heat transfer on the suction side weakens whilst it increases significantly on the pressure side. For the second channel, non-bleed condition leads to a more balanced heat transfer distribution between the upstream and downstream channel. Besides, after the bleed holes are blocked, heat transfer in the first bend region on the suction side declines sharply, while the opposite phenomenon occurs for the second bend region on the pressure side. In both bleed and non-bleed cases, the total pressure of six measurement positions decreases continuously along the channel at the same Reynolds number and it promotes for higher Reynolds number. Among all the measurement points, under the same flow rate condition, the highest speed occurs at Position 5, which also shows the maximum difference between the total and static pressures. When the bleed holes are blocked, the total pressure at each measurement position appears to increase.     

Equilibrium attitude and stability of a spacecraft on a stationary orbit around an asteroid

The stationary orbits around an asteroid, if exist, can be used for communication and navigation purposes just as around the Earth. The equilibrium attitude and stability of a rigid spacecraft on a stationary orbit around a uniformly-rotating asteroid are studied. The linearized equations of attitude motion are obtained under the small motion assumption. Then, the equilibrium attitude is determined in both cases of a general and a symmetrical spacecraft. Due to the higher-order inertia integrals of the spacecraft, the equilibrium attitude is slightly away from zero Euler angles. Then necessary conditions of stability of this conservative system are analyzed based on the linearized equations of motion. The effects of different parameters, including the harmonic coefficients C₂₀ and C₂₂ of the asteroid and higher-order inertia integrals of the spacecraft, on the stability are assessed and compared. Due to the significantly non-spherical shape and rapid rotation of the asteroid, the effects of the harmonic coefficients C₂₀ and C₂₂ are very significant, while effects of the third- and fourth-order inertia integrals of the spacecraft can be neglected. Considering a spacecraft on a stationary orbit around an example asteroid, we show that the classical stability domain predicted by the Beletskii–DeBra–Delp method on a circular orbit in a central gravity field is modified due to the non-spherical mass distribution of the asteroid. Our results are confirmed by a numerical simulation.     

航天器快捷交会技术分析

快捷交会对航天乘员与飞行任务都是非常有利的。快捷交会的基本原理是减少调相段飞行圈数;为此,需设计可完成一系列交会程序的较短的飞行时间,并获得与飞行时间相匹配的较小的初相角。快捷交会的关键技术包括提高入轨精度,增强追踪飞行器的自主计算与轨道控制能力,扩大地面站覆盖范围,以及对目标飞行器轨道的精准调整。     

A novel approach for predicting inlet pressure of aircraft hydraulic pumps under transient conditions

Cavitation caused by insufficient suction is a major factor that influences the life of aircraft pumps. Currently, pressurizing the tank can solve the cavitation problem under steady large-flow conditions. However, this method is not always effective under transient conditions (from zero flow to full flow in a very short time). Moreover, to apply and design other measures, such as a boost impeller, the suction dynamics during the transient period must be investigated. In this paper, a novel approach based on the pressure wave propagation theory is proposed for predicting the inlet pressure of an aircraft pump under transient conditions. First, a dynamic model of a typical aircraft pump is established in the form of differential equations. Then, the transient flow model of the inlet line is described using momentum and continuity equations, and the governing equations are discretized by the method of characteristics and the finite difference method. The simulated results are in good agreement with the results from verification tests. Further simulation analysis indicates that the wave velocity and transient time may influence the inlet and reservoir pressure as well as the size of the inlet line. Finally, solutions for upgrading the inlet pressure are discussed. These solutions provide guidelines for designing inlet installations.     

Optimum parameter matching obtained by experiments for coring drilling into lunar simulant

This paper performs a series of ground experiments (on the Earth) to obtain optimum parameter matching for a future core drill on the Moon. Three common stages, I, II, III, are recognized with respect to the cut per revolution (CPR), which is defined as the ratio between the feeding speed and the rotating speed. The optimum matching between the feeding speed and the rotating speed for drills locates at the boundary position between Stage II and Stage III, where the coring rate saturates and the weight on the bit and the driving torque are still low. Further data analysis of the ground experiments reveals that the optimum matching signifies a proportional relation between the maximum conveying rate (MCR) by the groove of the auger and its rotating speed. The kinetic analysis in an ideal condition without gravity, the friction from the auger groove and the pressure at the bit confirm a similar proportion. The correlation between the proportions needs further study to determine whether the optimum matching obtained on the ground can be directly applied to future drills on the Moon.     

基于模型的载人航天器研制方法研究与实践

载人航天器具有系统规模大、技术难度高、单件小批量、无法通过多次飞行持续完善设计、可靠性要求高等特点。当前载人航天器研制中仍存在着参数化和模型化程度不高、基于模型的系统综合仿真验证不足、研制各环节缺乏数字化集成等问题，传统基于文本的系统工程方法已无法满足研制需求，亟需采用基于模型的系统工程方法。本文针对载人航天器的研制现状和应用需求，提出了面向载人航天器全生命周期的模型体系，定义了需求模型、功能模型、产品模型、工程模型、制造模型、实做模型等六类模型，提出了基于模型的研制流程，包含系统设计闭环验证、产品设计闭环验证、实做产品闭环验证3个验证环节，并深入探索了各研制环节中不同模型间的传递与关联关系。以某型号载人航天器为应用基础，系统地验证了提出的方法。     

鱼游动涡结构PIV实验研究

鱼类逃逸和巡游已逐渐成为鱼类仿生推进水动力学领域的研究热点，其研究结果为水下航行器推进技术提供了很好的理论基础和指导意义。利用平面PIV技术测量了斑马鱼在水中游动时的尾迹流场，分析了不同游动状态下的鱼尾迹涡结构的变化规律；同时利用Tomo-PIV技术测量了曼龙鱼游尾迹三维流场，获得了涡环链结构。结果表明:不同游动状态下，鱼游尾迹表现出不同的流动结构和尾迹模式，对其进行研究有利于进一步揭示鱼游动的水动力学机理。     

Experimental investigation and correlation development of jet impingement heat transfer with two rows of aligned jet holes on an internal surface of a wing leading edge

Extensive experimental studies on the heat transfer characteristics of two rows of aligned jet holes impinging on a concave surface in a wing leading edge were conducted, where50000 Rej 90000, 1.74 H/d 27.5, 66° a 90°, and 13.2 r/d 42.03. The finding was that the heat transfer performance at the jet-impingement stagnation point with two rows of aligned jet holes was the same as that with a single row of jet holes or the middle row of three-row configurations when the circumferential angle of the two jet holes was larger than 30°. The attenuation coefficient distribution of the jet impingement heat transfer in the chordwise direction was so complicated that two zones were divided for a better analysis. It indicated that: the attenuation coefficient curve in the jet impingement zone exhibited an approximate upside-down bell shape with double peaks and a single valley; the attenuation coefficient curve in the non-jet impingement zone was like a half-bell shape, which was similar to that with three rows of aligned jet holes; the factors,including Rej, H/d and r/d, affected the attenuation coefficient value at the valley significantly.When r/d was increased from 30.75 to 42.03, the attenuation rates of attenuation coefficient increased only by 1.8%. Consequently, experimental data-based correlation equations of the Nusselt number for the heat transfer at the jet-impingement stagnation point and the distributionof the attenuation coefficient in the chordwise direction were acquired, which play an important role in designing the wing leading edge anti-icing system with two rows of aligned jet holes.     

Reconfigurable Flight Control Design for Combat Flying Wing with Multiple Control Surfaces

With control using redundant multiple control surface arrangement and large-deflection drag rudders,a combat flying wing has a higher probability for control surface failures.Therefore,its flight control system must be able to reconfigure after such failures.Considering three types of typical control surface failures(lock-in-place(LIP),loss-of-effectiveness(LOE) and float),flight control reconfiguration characteristic and capability of such aircraft types are analyzed.Because of the control surface redundancy,the aircraft using the dynamic inversion flight control law already has a control allocation block.In this paper,its flight control configuration during the above failures is achieved by modifying this block.It is shown that such a reconfigurable flight control design is valid,through numerical simulations of flight attitude control task.Results indicate that,in the circumstances of control surface failures with limited degree and the degradation of the flying quality level,a combat flying wing adopting this flight control reconfiguration approach based on control allocation could guarantee its flight safety and perform some flight combat missions.