Project title: Combined Effects of Roughness and Blowing in Ablatives

高速航天器在进入行星大气层时，由于动能转化为气体内部的内能，会产生令人难以置信的高气体温度。这些高温需要热保护系统（TPS），以减少航天器上的大热流，确保其在下降过程中的结构完整性。最常见的TPS类型是烧蚀热屏蔽，这是一层材料，当受到高热流时，会经历几个过程。这些过程包括产生多孔的烧焦纤维层和产生相对较冷的热解气体（热解是物质在高温下的分解）。烧焦的层导致非常粗糙的外表面，粗糙度的高度大大超过边界层厚度-紧邻航天器壁的薄层流体。这导致材料突出到高速流中，最终用于增加入射到航天器上的对流热通量（对流热通量是由于介质中存在温度梯度而以流体中的热的形式传递的能量）。为了抵消这种增加，将热解气体喷射到外表面上通过在周围流和材料之间产生冷缓冲层来减少入射热通量。这些现象和它们对入射到航天器壁上的热通量的总体增加已经通过大量的孤立实验得到了理解。但在重返大气层期间出现的现象则不能这样说。由于大尺度粗糙度和蒸发（将气体注入边界层作为冷却手段）的共同存在，几乎没有探索入射热通量的增加。对这些相互竞争的现象缺乏理解，导致模型的发展错误地对所产生的热通量进行了分类。我的研究，这福尔斯属于EPSRC工程研究领域，将有助于减少这种知识差距，通过实验，以帮助改善烧蚀模型的模拟，从而如何在实践中设计。主要目标是开发一个模型，从本质上讲，允许评估的总增加的热通量所造成的烧蚀热屏蔽。因此，模拟将更准确地封装烧蚀TPS在实践中的行为，从而更好地更有效地使用和设计它们。

High speed spacecrafts, when entering planetary atmospheres, produce incredibly high gas temperatures due to the conversion of kinetic energy into internal energy within the gas. These high temperatures create the need for thermal protection systems (TPS) put in place to reduce the large heat fluxes incident on the spacecraft, ensuring its structural integrity during descent. The most common type of TPS are ablatives heat shields, a layer of material that, when subject to high heat fluxes, undergoes several processes in response. These processes include the production of a porous layer of charred fibres and the generation of relatively cool pyrolysis gases (pyrolysis is the decomposition of a substance as a consequence of high temperatures). The charred layer results in a very rough external surface, with a roughness height that greatly exceeds the boundary layer thickness - the thin layer of fluid in the immediate vicinity of the spacecraft walls. This results in material protruding into the high-speed flow ultimately serving to increase the convective heat flux incident on the spacecraft (a convective heat flux is energy transferred in the form of heat in a fluid due to the presence of a temperature gradient in the medium). Counteracting this increase, the injection of the pyrolysis gases onto the external surface reduces the incident heat flux through the generation of a cold buffer layer between the surrounding flow and the material. These phenomena and their overall augmentation to the heat flux that would be incident on the spacecraft walls have been understood through a plethora of experiments in isolation. The same cannot be said for the phenomena as they occur during re-entry. There as been little exploration into the augmentation of the incident heat flux due to the combined presence of roughness on large scales and transpiration (the injection of gas into the boundary layer as a means of cooling). This lack of understanding of these competing phenomena in conjunction results in the development of models that incorrectly categorise the resultant heat flux. My research, which falls within the EPSRC engineering research area, will contribute to the lessening of this knowledge gap through experimentation to aid the improvement of ablative models in simulation and thus how they are designed in practise. The key goal is to develop a model that, in essence, allows the evaluation of the total augmentation to the heat-flux caused by the ablative heat shields. As a result, simulations will more accurately encapsulate the behaviour of ablative TPS in practise leading to better more efficient use and design of them.