Project title: Combined Effects of Roughness and Blowing in Ablatives

高速航天器进入行星大气层时，由于动能转化为气体内部能量，会产生极高的气体温度。在这种高温下，需要安装热保护系统（TPS），以减少航天器上的大热通量，确保其在下降过程中的结构完整性。最常见的TPS类型是烧蚀隔热板，这是一层材料，当受到高热流通量时，会经历几个过程作为响应。这些过程包括产生多孔的烧焦纤维层和产生相对凉爽的热解气体（热解是高温下物质的分解）。烧焦层导致了非常粗糙的外表面，其粗糙度高度大大超过了附面层厚度——即航天器壁附近的薄层流体。这导致物质凸出到高速气流中，最终增加了入射到航天器上的对流热流（对流热流是由于介质中存在温度梯度而以热形式在流体中传递的能量）。通过在周围气流和物料之间形成冷缓冲层，将热解气体注入外表面来抵消这种增加，从而减少了入射热通量。通过大量的孤立实验，人们已经了解了这些现象以及它们对入射到航天器壁上的热通量的总体增强。再入大气层时发生的现象就不能这么说了。由于大尺度上的粗糙度和蒸腾作用（将气体注入边界层作为冷却手段）的共同存在，对入射热通量的增加进行了很少的探索。由于缺乏对这些相互竞争的现象的理解，导致了模型的发展，这些模型错误地对所得热通量进行了分类。我的研究属于EPSRC工程研究领域，将有助于通过实验来减少这种知识差距，以帮助改进模拟中的烧蚀模型，从而在实践中如何设计它们。关键目标是建立一个模型，从本质上讲，该模型允许评估由烧蚀隔热板引起的热通量的总增量。因此，模拟将更准确地概括烧蚀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.