Hypervelocity impact induced shock acoustic emission waves for quantitative damage evaluation using in situ miniaturized piezoelectric sensor network

Manmade debris and natural meteoroids, travelling in the Low Earth Orbit at a speed of several kilometers per second, pose a severe safety concern to the spacecraft in service through the HyperVelocity Impact(HVI). To address this issue, an investigation of shock Acoustic Emission(AE) waves induced by HVI to a downscaled two-layer Whipple shielding structure is performed,to realize a quantitative damage evaluation. Firstly a hybrid numerical model integrating smoothparticle hydrodynamics and finite element is built to obtain the wave response. The projectiles, with various impact velocities and directions, are modelled to impact the shielding structure with different thicknesses. Then experimental validation is carried out with built-in miniaturized piezoelectric sensors to in situ sense the HVI-induced AE waves. A quantitative agreement is obtained between numerical and experimental results, demonstrating the correctness of the hybrid model and facilitating the explanation of obtained AE signals in experiment. Based on the understanding of HVI-induced wave components, assessment of the damage severity, i.e., whether the outer shielding layer is perforated or not, is performed using the energy ratio between the regions of ‘‘high frequency" and ‘‘low frequency" in the acquired AE signals. Lastly, the direct-arrival fundamentalsymmetric wave mode is isolated from each sensing signal to be input into an enhanced delay-andsum algorithm, which visualizes HVI spots accurately and instantaneously with different sensor network configuration. All these works demonstrate the potential of quantitative, in situ, and real time HVI monitoring using miniaturized piezoelectric sensor network.     

Effects of Mach numbers on Magnus induced surface pressure

Experimental and numerical methods were used to investigate the Magnus phenomena over a spinning projectile. The pressure force acting on the surface of a spinning projectile was measured for various cases by employing a relatively novel experimental technique. A set of miniature pressure sensors along with a data acquisition board, battery and storage memory were placed inside a spinning model and the surface pressure were obtained through a remotely controlled system. Circumferential pressures of the model for both rotational and static conditions were obtained at two different free stream Mach numbers of 0.4 and 0.8 and at different angles of attack. The results showed the ability of this new test method to measure the very small Magnus force via surface pressures over the projectile. The results provide a deep insight into the flow structure and illustrate changes in the cross-flow separation locations as a result of rotation. Similar results were obtained by the numerical simulations and were compared with the experimental data.     

Numerical investigation on flow nonuniformity-induced hysteresis in scramjet isolator

Numerical simulation and theoretical analysis were conducted to study the hysteresis inside scramjet isolator during the reciprocating process of back pressure variation. It is revealed that only a regular reflection is theoretically possible for two leading shocks when the inflow Mach number is greater than 2.0, and no hysteresis can occur in the transition between shock reflection types. Nevertheless, wall suction, gas injection, and background waves cause non-uniformity of the incoming flow and would make hysteresis possible. Besides the classical hysteresis in the transition between shock reflection, new kinds of hysteresis were found in both the deflection angle of separated boundary layer and the location of the shock train. Moreover, the occurrence of hysteresis in the deflection angle of the separated boundary layer is accompanied with the shock reflection hysteresis. In the case with background waves or gas injection, hysteresis in the starting position of leading shock was observed too. As back pressure decreases, the leading shock does not follow the same path as that as the back pressure increases, and it is anchored at the location where the background shock or the injection interacts with the leading shock. It is inferred that, if two strong adverse pressure gradient regions move towards and interact with each other, hysteresis will take place when they start to separate.     

A comprehensive survey on the methods of angle of attack measurement and estimation in UAVs

Angle of Attack (AOA) is a crucial parameter which directly affects the aerodynamic forces of an aircraft. The measurement of AOA is required to ensure a safe flight within its designed flight envelop. This paper intends to summarise a comprehensive survey on the measurement techniques and estimation methods for AOA, specifically in Unmanned Aerial Vehicle (UAV) applications. In the case of UAVs, weight constraint plays a major role as far as sensor suites are concerned. This results in selecting a suitable estimation method to extract AOA using the available data from the autopilot. The most feasible and widely employed AOA measurement technique is by using the Multi-Hole Probes (MHPs). The MHP measures the AOA regarding the pressure variations between the ports. Due to the importance of MHP in AOA measurement, the calibration methods for the MHP are also included in this paper. This paper discusses the AOA measurement using virtual AOA sensors, their importance and the operation.     

静压气浮轴承工程设计方法研究

静压气浮轴承是惯性器件测试平台回转支撑轴系的关键部件，其承载能力和回转精度对惯性器件测试的准确性有着重要的影响，目前，静压气浮轴承设计、计算过程复杂，为提高设计效率，简化设计方法，在满足静压气浮轴承性能的条件下，推导了静压气浮轴承工程计算公式，在此基础上，设计了一款双向止推空气轴承，确定了几何参数，计算了空气静压径向轴承、空气静压止推轴承的轴承承载力、刚度、耗气量及摩擦力矩，并与有限元仿真结果进行了对比分析，验证了该设计方法的有效性。     

沸腾换热在水卡式大热流传感器研制中的应用

为了实现MW级热流环境的长时间稳态测量,研制了一种新型结构的水卡式大热流传感器。该传感器采用U型水冷通道结构,基于沸腾换热理论进行换热设计,设计量程5~15 MW/m~2。在高温超声速燃气流试验台上进行了热流测试试验,实际测试量程3~25 MW/m~2。沸腾换热理论的应用可以在较小的水流速下实现MW级热防护,大大降低了设计难度。新型结构传感器相对传统的中心冲击结构换热效率更高,可实现的量程范围更大。     

Fast spacecraft solar radiation pressure modeling by ray tracing on graphics processing unit

A novel method is presented to evaluate on the graphics processing unit (GPU) the force and torque on a spacecraft due to solar radiation pressure. The method employs efficient ray tracing techniques, developed in the graphics rendering discipline, to resolve spacecraft self-shadowing and reflections at faster than real-time computation speed. The primary algorithmic components of the ray tracing process which contribute to the method’s computational efficiency are described. These components include two-level bounding volume hierarchy acceleration data structures, fast ray to bounding box intersection testing using the slab intersection algorithm and fast triangle intersection testing using the Möller-Trumbore algorithm. Spacecraft material optical properties are represented as a combination of Lambertian diffuse and ideal specular reflections. Both diffuse and specular ray-surface interactions are modeled. The approach is implemented using C++ and OpenCL and executed on a consumer grade GPU. Model validation is presented comparing ray traced force and torque values to the same quantities produce by a faceted analytic model. Numerical results illustrate the impact of self-shadowing on the force and torque calculation, and demonstrate the fast computational speed that is enabled with this implementation.     

主流攻角对双射流孔气膜冷却特性的影响

应用压力敏感漆技术,在平板上测量了不同主流攻角(i=-30°,-20°,-10°,0°,10°,20°,30°)下双射流孔的气膜冷却效率,并利用计算流体动力学(CFD)计算得到的流场对气膜冷却效率的规律进行了解析。所研究的双射流孔结构的孔间无量纲横向距离为0.5,孔间无量纲流向距离为3;射流与主流密度比为1.0,吹风比分别为0.5、1.0、1.5、2.0。结果表明小的主流攻角(i=-10°,0°,10°)下,流场中存在反肾型涡对或挤压作用,气膜层与壁面贴附良好,气膜冷却效率最高;大正值攻角(i=20°,30°)下,虽然气膜覆盖面积大,但反肾型涡对退化,气膜冷却效率下降;大负值攻角(i=-20°,-30°)下,流场中有肾型涡对,且气膜横向覆盖受限,气膜冷却效率最低。     

盘腔进气位置对径向预旋系统的影响

为减少径向预旋系统的流动损失,运用数值模拟方法对不同盘腔进气位置的径向预旋系统进行分析,结果表明:随着盘腔进气径向位置的增加,预旋喷嘴出口气流旋流比随之逐渐减小,径向预旋系统的温降系数及总压损失系数均随之逐渐增大。当旋转雷诺数等于7.9×106,盘腔进气位置由低位向高位变化时温降系数最大可增加525%,同时总压损失系数增加3.93%。径向预旋系统内比熵增主要发生在预旋喷嘴和共转腔,约占系统总体比熵增的80%。随着盘腔进气径向位置的增加,径向预旋系统总体比熵增降低,预旋喷嘴比熵增占比逐渐增大,共转腔比熵增占比逐渐减小。     

极低压下两相流挤压油膜阻尼器油膜参数特性

针对极低供油压力工况开展实验研究,以考察气液两相流对挤压油膜阻尼器(SFD)油膜参数特性的影响。结果表明:当SFD入口气体体积分数小于0.9时,油膜阻尼随着入口气体体积分数的增加而减小,直至气体体积分数增大到0.9时,油膜阻尼是纯油状态时阻尼的60%;当气体体积分数大于0.9时,油膜阻尼大幅减小至几乎可以忽略不计。现有理论模型并不适用于极低供油压力工况。基于实验结果,找到了最符合SFD两相流动的等效黏度模型,其理论预测的油膜阻尼与不同供油压力下的实验数据吻合较好,为SFD两相流研究的模型选择提供了依据。