深空撞击载荷总体技术分析与效能仿真

以近地小行星2016HO3为深空撞击目标开展科学探测，从爆炸成型弹丸技术、安全可靠爆炸技术、撞击载荷分离技术、撞击效能仿真技术等7个方面系统详细地进行了基于聚能爆炸成型弹丸的撞击载荷技术体系解析与技术内涵阐述，给出了深空撞击载荷初步总体设计方案；通过弹丸撞击靶板的数值模拟仿真，得到了撞击速度在0.2～0.4cm/μs变化范围下产生32～47cm大小的撞击坑坑径结果，得出了撞击坑直径随弹丸不同撞击速度、靶材不同密度、不同强度和不同厚度的变化规律，为撞击载荷总体优化设计提供了撞击效能仿真技术支撑。     

柱状弹丸撞击防护屏形成碎片云材料状态特性研究

微流星体及空间碎片的高速撞击威胁着长寿命、大尺寸航天器的安全运行，导致其严重的损伤和灾难性的失效。为精确估计微流星体及空间碎片高速撞击防护屏所产生碎片云对舱壁的损伤，必须确定碎片云中三种状态材料的特性，建立了碎片云特性分析模型，分别计算了柱状弹丸撞击防护屏所产生碎片云以及碎片云中弹丸和防护屏材料三种状态物质的质量分布。通过计算分析可见，弹丸以不同速度撞击防护屏所产生碎片云三种状态物质的质量分布是不同的，速度增大，液化和气化增强，对靶件的损伤小。而在速度小于7km/s时，碎片云以固体碎片的形式存在，对靶件的损伤大。     

基于缩放实验的Whipple结构撞击特性

针对航天器空间碎片防护问题,基于缩放实验方法,开展了7 km/s以上超高速碰撞仿真研究.建立了单板和Whipple防护结构的仿真模型,并对铝-铝撞击问题和镉-镉撞击问题进行了多工况仿真.通过实验结果与数值仿真的对比,表明了数值仿真技术的正确性,并从仿真角度验证了缩放实验方法的有效性.对缩放实验方法的适用性进行了仿真验证,结果表明该方法对弹丸形状适用性较好,对3～4 km/s以上撞击速度的适用性较好,但对Whipple防护结构后板存在一定误差.分析了Whipple结构后板的失效模式,提出了失效模式的不连续性导致了缩放实验方法的误差.最后通过数值仿真计算了Whipple结构7 km/s以上弹道极限特性,提出了失效模式的不连续性造成了在该速度段弹道极限曲线的分叉现象.     

空间碎片的防护、监测和建模

正空间碎片与航天器的平均撞击速度为10公里/秒,这么高的撞击速度,现有材料难以"扛得住"。那么如何对这种撞击进行防护呢?空间碎片的防护实验证明,超高速弹丸(碎片)与薄靶撞击过程中,会发生破碎、熔化、气化甚至等离子体化等,形成高速运动的物质云团,称为碎片云。弹丸和薄靶     

卫星高压气瓶的超高速撞击试验

微流星体及空间碎片超高速撞击对在轨航天器构成了严重威胁,星上压力容器受空间碎片撞击后所产生的威胁是十分严重的,可能导致航天器发生灾难性失效,过早结束其使命。文章通过星上常用气瓶的超高速撞击试验,获取了不同弹丸撞击参数下气瓶器壁的通孔孔径,得到了在弹丸撞击速度为(6.5±0.3)km/s、无防护情况下气瓶器壁的弹道极限,并分析了导致充压气瓶灾难性失效的弹丸直径范围;通过对试验数据拟合,初步建立了弹丸正撞击速度为(6.5±0.3)km/s、无防护情况下气瓶器壁的通孔孔径预测公式,为航天器遭遇空间碎片撞击的风险评估及防护措施制定提供依据。     

高速撞击充气压力容器前壁损伤数值模拟

针对空间碎片超高速撞击充气压力容器前壁损伤问题,应用非线性动力学分析软件AUTODYN采用拉格朗日方法对球形弹丸撞击球形压力容器前壁穿孔进行了数值模拟研究。在建模过程中通过对容器壁内侧施加压力边界条件来模拟由于内充气体的作用在容器壁内产生的应力场,并通过与试验结果的比较验证了数值模拟方法的有效性。在此基础上针对容器的内充气体压力、球形弹丸直径及撞击速度对充气压力容器前壁穿孔的影响进行了研究。结果表明:在一定的气体压力下,气体压力对压力容器前壁穿孔直径与穿孔形态的影响可以忽略不计;而撞击速度及弹丸直径对穿孔直径及穿孔形态有着较大的影响,当撞击速度大于3km/s时,撞击穿孔边缘开始有裂纹产生,并且穿孔直径与裂纹直径随着弹丸直径及撞击速度的增加而增大。利用压力容器前壁穿孔的数值模拟结果进行计算可以得出当容器受到撞击速度大于3km/s的弹丸撞击后比撞击速度不大于3km/s时更易发生破坏。     

弹丸撞击防护屏形成碎片云速度特性研究

微流星及空间碎片的高速撞击威胁着长寿命，大尺寸航天器的安全运行，导致其严重的损伤和灾难性的失效，为精确估计微流星及空间碎片主速撞击防护屏产生的碎片对舱壁的损伤，必须确定碎片云速度特性。文章在冲量和能量守恒的基础上，建立了碎片速度性分析模型，研究了碎片云的速度特性，得到了碎片云材料传播及碎片云喷射角随弹丸撞击速度的变化规律。     

基于Adaboost的填充式防护结构超高速撞击损伤预测

填充式防护结构的显式弹道极限方程在对弹丸进行超高速撞击损伤预测时,由于填充材料、填充方式的不同,会导致预测结果与实测数据存在一定偏差。对此,采用机器学习方式将该问题转化为二分类问题,以碰撞过程中的弹丸撞击参数、防护结构参数作为分类特征,构建了基于Adaboost的填充式防护结构超高速撞击损伤预测模型。该模型以分类回归树（CART）作为弱分类器,通过对一系列弱分类器的加权组合生成强分类器,并通过对训练样本的循环使用,实现了小样本集下的撞击损伤预测。实验结果表明,建立的Adaboost预测模型对填充式防护结构的超高速撞击损伤具有良好的预测效果,总体预测率与安全预测率相比于NASA的弹道极限方程均提高了14.3%,具有更强的通用性。通过不同训练样本规模下的交叉检验,证明了该模型具有良好的鲁棒性与准确性。      

美国国防部研制轨道枪

美国国防部正着手制定一项研制适用于美国航天飞机对抗苏联猎潜艇卫星的轨道枪系统计划。这种轨道枪将能够发射弹丸,其速度超过10公里/秒(6.2哩/秒),利用动能去消灭它的攻击目标。国防高级研究计划局(DARPA)和空军正在用从电磁轨道枪发射的小弹丸在真空容器里以8.6公里/秒的空间速度进行实验。在这些实验里,电能转换到磁压力推进装     

双柔性杆多次撞击力分析

柔性构件的多次弹性撞击现象在工程中普遍存在,其中撞击力的求解一直是关键问题.利用瞬态波函数特征值展开法,给出多次撞击过程和分离过程的杆中瞬态波传播的理论解,提出计算双柔性杆多次撞击力的瞬态波效应法,该方法避免求解具有未知奇异载荷项的强非线性方程.实现了对多次撞击问题的精确的数值分析,研究了初始撞击速度对撞击力和撞击发生时间的影响,表明瞬态波效应法可以计算多次撞击力时间历程.初始撞击速度越大,撞击力幅值越大,而且撞击力的峰值在前三次撞击过程中是逐次增加的,后趋于稳定,并可以研究"次撞击"现象.为进一步研究复杂柔性结构系统多次撞击问题以及重复撞击引起的长期系统动力学行为提供了可借鉴的有效研究方法.      

超高速撞击航天器二次碎片云能量特性分析

为了评估空间碎片超高速撞击航天器的碎片云破坏能力,挖掘超高速撞击数值模 拟结果数据的应用价值,基于9.53 mm铝球以6.64 km/s速度对2.2 mm铝靶撞 击的Ls-Dyna/SPH(Smoothed Particle Hydrodynamic)数值模拟研究结果,对靶后碎片云的 粒子动能进行求和统计,建立了碎片云比动能概念和函数形式;碎片云比动能综合考虑了靶 后所有碎片云粒子的动能,反映了一定距离处垂直于撞击方向平面上单位面积上的碎片云粒 子所蕴含的撞击能量;应用碎片云比动能概念,揭示出随着演化距离的增加,碎片云能量的 衰减规律;通过不同速度条件下的SPH计算,得到了碎片云的比动能函数的曲线形式随撞击 速度的变化规律;最后对采用2种材料模型进行数值模拟所对应的结果误差进行碎片云比动 能函数的曲线比较,反映出数值模拟中不同材料模型引起的差异.      

空间碎片超高速撞击动力学建模与数值仿真技术

阐明了空间碎片超高速撞击数值仿真技术研究的目的、意义和国内外发展状况 ;重点论述了空间碎片超高速撞击数值仿真技术的主要研究内容、技术指标和具体实施途径 ,从而为研究的深入开展提供了技术依据和指导原则     

Micro-meteoroids and space debris impact risk assessment for the ConeXpress satellite using ESABASE2/Debris

Micro-meteoroid and space debris impact risk assessments are performed to investigate the risk from hypervelocity impacts to sensitive spacecraft sub-systems. For these analyses, ESA’s impact risk assessment tool ESABASE2/Debris is used. This software tool combines micro-particle environment models, damage equations for different shielding designs and satellite geometry models to perform a detailed 3D micro-particle impact risk assessment. This paper concentrates on the impact risk for exposed pressurized tanks. Pressure vessels are especially susceptible to hypervelocity impacts when no protection is available from the satellite itself. Even small particles in the mm size range can lead to a fatal burst or rupture of a tank when impacting with a typical collision velocity of 10–20 km/s. For any space mission it has to be assured that the impact risk is properly considered and kept within acceptable limits. The ConeXpress satellite mission is analysed as example. ConeXpress is a planned service spacecraft, intended to extend the lifetime of telecommunication spacecraft in the geostationary orbit. The unprotected tanks of ConeXpress are identified as having a high failure risk from hypervelocity impacts, mainly caused by micro-meteoroids. Options are studied to enhance the impact protection. It is demonstrated that even a thin additional protective layer spaced several cm from the tank would act as part of a double wall (Whipple) shield and greatly reduce the impact risk. In case of ConeXpress with 12 years mission duration the risk of impact related failure of a tank can be reduced from almost 39% for an unprotected tank facing in flight direction to below 0.1% for a tank protected by a properly designed Whipple shield.     

Analysis of space debris impact source localization based on PVDF piezoelectric film

As the pace of human exploration and utilization of space continues to accelerate, space debris gradually becomes an inevitable problem affecting and threatening human space activities. When space debris strikes the spacecraft bulkhead, determining the impact source location timely and accurately is the foundation of the repair damage, and is also of great importance for the safety of astronauts' life. This paper analyzed the wave propagation law in thin plates, established a lightweight sensor array using PVDF (Polyvinylidene fluoride) circular thin-film sensors, and used a two-stage light-gas gun loading system to conduct hypervelocity collision localization experiments on impacting 2A12 aluminum plates to study the effects of sensor array radius and sensor size on localization results. The results show that the smaller the radius of the PVDF sensor array is, the more accurate the positioning result is under the premise of the same size of the PVDF circular film sensor array. On the premise of the same PVDF sensor array arrangement, the larger the PVDF circular membrane sensor is, the more accurate the positioning result is. ABAQUS finite element software is used to study the stress wave propagation of aluminum ball impacting aluminum plate at high speed, simulating space debris impacting spacecraft. The stress waveform obtained from the simulation is in good agreement with the experiment, which shows the accuracy of the numerical simulation method.     

Finite element reconstruction approach for on-orbit spacecraft breakup dynamics simulation and fragment analysis

Breakup model is the key area of space debris environment modeling. NASA standard breakup model is currently the most widely used for general-purpose. It is a statistical model found based on space surveillance data and a few ground-based test data. NASA model takes the mass, impact velocity magnitude for input and provides the fragment size, area-to-mass ratio, velocity magnitude distributions for output. A more precise approach for spacecraft disintegration fragment analysis is presented in this paper. This approach is based on hypervelocity impact dynamics and takes the shape, material, internal structure and impact location etc. of spacecraft and impactor, which might greatly affect the fragment distribution, into consideration. The approach is a combination of finite element and particle methods, entitled finite element reconstruction (FER). By reconstructing elements from the particle debris cloud, reliable individual fragments are identified. Fragment distribution is generated with undirected graph conversion and connected component analysis. Ground-based test from literature is introduced for verification. In the simulation satellite targets and impactors are modeled in detail including the shape, material, internal structure and so on. FER output includes the total number of fragments and the mass, size and velocity vector of each fragment. The reported fragment distribution of FER shows good agreement with the test, and has good accuracy for small fragments.     

Progress in Space Debris Research

During recent years, A de-orbit disposal of SinoSat 2 satellite and the depletion of the residual propellant after SC/LV separation for all LM-4 series launch vehicles were carried out. Stuffed Whipple Shields based on hypervelocity impact particles were developed. Routine observation and collision avoidance were performed. The main progress in space debris research will be introduced from three aspects: mitigation, spacecraft protection, observation and collision avoidance.      

Hypervelocity impact on the GIOTTO Halley Mission dust shield: Momentum exchange and measurement

The two layer dust shield on the GIOTTO Halley Mission is constructed in a meteoroid bumper configuration. The dust shield is instrumented so that parameters associated with the hypervelocity collision of cometary particles on the exposed surface can be determined. A multisensor detector array provides simultaneous sensing of the momentum exchange of particles impacting and subsequently penetrating the outer layer of the dust shield. Current knowledge of momentum exchange during hypervelocity impact relative to the GIOTTO Halley Mission and the dust shield experiment is reviewed. The sensors used for determination of momentum exchange exhibit a functional dependence on projectile velocity leading to an enhancement of the sensor signal as the relative impact velocity increases. The GIOTTO Mission provides a very unique opportunity to obtain hypervelocity momentum exchange information at a known impact velocity. Therefore, with the dust experiment, a determination of the velocity index for both momentum and multilayered penetration sensor is possible. Results of analysis of analytical and laboratory studies indicate that the velocity index for hypervelocity impact is approximately 2.0 at the 68 km/sec encounter impact velocity of the GIOTTO Mission. A clear determination of the size and mass distribution of the cometary dust near the comet will be possible from the in-situ measurement of the DIDSY GIOTTO experiment.