离子推力器羽流热效应仿真分析

离子推力器工作时向外喷出的羽流与航天器表面碰撞,会引起敏感材料热变形等热效应,严重时会导致航天任务失败。针对兰州空间技术物理研究所研制的LIPS-200型离子推力器羽流热效应进行了仿真分析。仿真中,使用粒子网格（PIC）方法处理等离子体运动,使用直接模拟蒙特卡罗（DSMC）方法处理粒子间碰撞,使用Maxwell模型处理粒子与壁面的能量交换,对电推进羽流热效应测量中的部分测点进行了数值模拟。结果表明,仿真结果与实验数据符合较好,离子推力器出口轴线上滞止热流仿真值与实验测量值误差小于17.0%。此外,热流计对流场的影响主要集中在热流计附近0.1 m范围内,对整体流场影响较小。      

基于双向拉伸的热环境铝合金性能获取和分析

板材成形加工时通常承受复杂载荷,一般采用单拉试验获取材料性能,由于材料变形时仅承受单向载荷,与实际情况差距较大。为获取更加真实的复杂加载时材料性能,通过十字形试件双向拉伸试验,研究了热环境双向变比例加载时AA6016铝合金材料力学性能和变形行为,包括优化设计十字形试件、相关试验方法和设备以及结果分析等。在25、150和250℃温度下进行了拉伸速率比例为1:1、3:2、2:3、1:3和3:1的双向拉伸试验和单向拉伸试验,得到了不同拉伸速率比例和温度下的应力应变关系、屈服规律和各向异性,建立了屈服准则,并且通过与试验结果对比,讨论分析了几个典型屈服准则及其适用性。      

火星表面热环境对航天器热控影响分析

火星大气对太阳辐射产生吸收和散射作用,同时还将与火星表面航天器发生对流换热.热设计时难以直接评估对流、辐射和导热三种换热对航天器的影响,从而确定主要的控温途径.在调研火星表面辐射、大气等热环境的基础上,从线性化传热系数和对流辐射比的角度对比分析了辐射、对流和导热对航天器的影响.器表辐射传热系数随光学属性和温度的变化范围...     

时变热辐射环境下高温合金蜂窝板三维热变形测量

轻质、高强和隔热性能优良的蜂窝合金板结构已广泛用于航空航天领域,其在模拟瞬态气动热环境下的热变形是高速飞行器热防护结构设计的重要参数之一.首先,用自行研制的红外辐射瞬态气动热实验模拟系统模拟与其服役环境类似的时变热辐射环境,用新型"主动成像"三维数字图像相关(3D-DIC)测量方法对高温合金蜂窝板结构试件在时变热辐射环境下不同时刻的三维热变形进行了测量.其次,为保证三维数字图像相关测量方法能有效实施,提出一种制作稳定的大面积高温散斑新方法,该方法制作的高温散斑能在整个实验过程中保持稳定,可作为高温变形的有效载体.最后,用Hoff等效刚度理论计算高温合金蜂窝板在稳态时的最大翘曲位移.研究结果表明:210 mm×210 mm的高温合金蜂窝板在单侧面辐射加热条件下其面内变形为均匀热变形,而离面变形为轴对称的翘曲变形,在900℃时其最大离面翘曲位移约为1.6 mm;Hoff等效刚度理论计算结果与实验结果相吻合.      

DES方法对返回舱气动热环境数值模拟

脱体涡模拟方法(DES,Detached Eddy Simulation)由于结合了湍流模型(RNAS,Reynolds Averaged Navier-stokes)和大涡模拟(LES,Large Eddy Simulation)两者各自优势,是模拟非定常大分离流动的有效方法.采用基于SA (Spalart-Allmaras)湍流模型的DES方法,对高超声速返回舱外形进行数值模拟,计算所得的返回舱表面压力及热流分布与实验结果吻合,证实了DES方法的优越性.最后,在DES方法中加入可压缩修正,结果证实在高超声速流动中,可压缩修正方法在一定程度上能够提高原始模型的预测性能.     

DES方法对返回舱气动热环境数值模拟

脱体涡模拟方法(DES,Detached Eddy Simulation)由于结合了湍流模型(RNAS,Reynolds Averaged Navier-Stokes)和大涡模拟(LES,Large Eddy Simulation)两者各自优势,是模拟非定常大分离流动的有效方法.采用基于SA(Spalart-Allmaras)湍流模型的DES方法,对高超声速返回舱外形进行数值模拟,计算所得的返回舱表面压力及热流分布与实验结果吻合,证实了DES方法的优越性.最后,在DES方法中加入可压缩修正,结果证实在高超声速流动中,可压缩修正方法在一定程度上能够提高原始模型的预测性能.     

Degradation of thermal control materials under a simulated radiative space environment

A spacecraft with a passive thermal control system utilizes various thermal control materials to maintain temperatures within safe operating limits. Materials used for spacecraft applications are exposed to harsh space environments such as ultraviolet (UV) and particle (electron, proton) irradiation and atomic oxygen (AO), undergo physical damage and thermal degradation, which must be considered for spacecraft thermal design optimization and cost effectiveness. This paper describes the effect of synergistic radiation on some of the important thermal control materials to verify the assumptions of beginning-of-life (BOL) and end-of-life (EOL) properties. Studies on the degradation in the optical properties (solar absorptance and infrared emittance) of some important thermal control materials exposed to simulated radiative geostationary space environment are discussed. The current studies are purely related to the influence of radiation on the degradation of the materials; other environmental aspects (e.g., thermal cycling) are not discussed. The thermal control materials investigated herein include different kind of second-surface mirrors, white anodizing, white paints, black paints, multilayer insulation materials, varnish coated aluminized polyimide, germanium coated polyimide, polyether ether ketone (PEEK) and poly tetra fluoro ethylene (PTFE). For this purpose, a test in the constant vacuum was performed reproducing a three year radiative space environment exposure, including ultraviolet and charged particle effects on North/South panels of a geostationary three-axis stabilized spacecraft. Reflectance spectra were measured in situ in the solar range (250–2500 nm) and the corresponding solar absorptance values were calculated. The test methodology and the degradations of the materials are discussed. The most important degradations among the low solar absorptance materials were found in the white paints whereas the rigid optical solar reflectors remained quite stable. Among the high solar absorptance elements, as such the change in the solar absorptance was very low, in particular the germanium coated polyimide was found highly stable.     

海洋环境下三电极的电偶腐蚀仿真

为研究三电极的电偶腐蚀行为,测量了CF8611/AC531复合材料（CFRP）、7B04-T74铝合金（7B04）和镀锌30CrMnSiA钢（GSB）的极化曲线;开展了搭接件在模拟海洋环境下的全浸试验;设计了圆形三电极,推导了稳态腐蚀场和参数化扫描方程,建立了三电极和搭接件的电偶腐蚀模型。结果表明:稳态腐蚀场中的电势分布符合Laplace方程;电位最高的CFRP为阴极,最低的GSB为阳极,中间的7B04阴/阳极角色会随某一电极面积变化而转变,给出了转变的临界面积比,各电极表面电偶电流服从指数分布,相关系数近于1,拟合精度高;在搭接件中,搭接区电位和电流密度最高,并向两端对称递减,7B04和GSB均为阳极,电流密度分别提高约210倍和328倍,电偶腐蚀效应显著;搭接区7B04板全面腐蚀,厚度损失约1.011%;仿真所得点蚀敏感区宽度范围为3.9~7.6mm,实测所得宽度范围为4.667~8.872mm,二者范围、形状及变化规律吻合较好,表明模型有效、可靠。      

Feasibility analysis of solar thermal propulsion system with thermal energy storage

Possessing relatively high specific impulse and moderate thrust levels, solar thermal propulsion (STP) is a promising candidate in spacecraft propulsion system. However, the traditional solar thermal propulsion system suffers from thrust failure in the shadow area, which seriously affects its applicability. In this paper, we investigate feasibility of regenerative solar thermal propulsion system (RSTP) incorporating thermal energy storage, which can effectively overcome unmatched synchronous working time and illumination time. A numerical model for RSTP considering the whole energy transfer process from light concentrating, heat storage, to thrust generation is built, which is verified by experiment measurements with relative errors less than 15 %. The result shows that the maximum time to complete heat storage is about 4000 s, which is within the illumination time for low Earth orbit. In the solar eclipse region, the thrust (F_t) and the specific impulse (I_sp) of the system increase with the propellant flow rate, which can reach about 2 N and 690 s, respectively. What’s more, the system can operate for around 100 s continuously at the maximum thrust in the shadow area. This work provides alternative approaches for microsatellite propulsion with high specific impulse, high thrust, and continuous operation despite presence of solar eclipse.     

环境试验设备的状态空间法仿真设计

对于环境试验设备的设计,传统的工程设计方法是从指标出发,用半经验、稳态的方法设计系统,按照围护结构稳态放热量和试品冷透来计算制冷系统容量,这种稳态的计算结果通常比实际需要大的多.采用基于状态空间描述的分布参数模型用于系统动态仿真,可以允许对许多先不必要考虑的环节简化处理,对主要的负荷如围护结构、被试品等做精确的处理,可以较准确的得到总的制冷量需求,验证控制策略等,对于初期方案设计中系统制冷、加热设备容量的准确计算、不同方案的验证、控制方案的制定具有明显的意义.