Global Measurements of Inlet/Isolator Dynamics and Optimization of Unstart Detection

Dual-mode scramjets are a promising technology to expand the capabilities of hypersonic vehicles. Streamtraced inlets are an attractive inlet concept for high performance dual-mode scramjets. These inlets generate complex shock-wave/boundary-layer interactions, which have not been widely studied in the literature, especially with high-resolution measurements. Scramjet unstart is a phenomenon that results in a severe reduction in inlet mass capture and occurs during operation at off-design conditions. Unstart can occur in milliseconds, and the flow physics governing it are extremely complex, involving numerous shock-wave/boundary-layer interactions. Improved prediction and detection of unstart is required to inform proper engine design and allow for effective flow control to avoid unstart. In this study, global surface measurements are applied to axisymmetric and streamtraced inlet/isolators to better measure shock-wave/boundary layer interactions and unstart. Tests were conducted in the AFOSR--Notre Dame Large Mach-6 Quiet Tunnel under Mach 5.7 flow for a variety of unit Reynolds numbers. Transverse jet injection of air at the aft portion of the isolator induced unstart of the inlet under cold-flow conditions. Fast pressure-sensitive paint (PSP) was used to measure the time-resolved pressure field on the inside of the isolator. The PSP was temperature compensated using temperature-sensitive paint measurements. Background-oriented schlieren was utilized for visualizing the inlet shock structure during unstart. The utility of PSP for investigating the dynamics of internal SWBLI was demonstrated using a variety of data-analysis techniques. The mean shock locations in the isolator were determined to within 0.1~mm. Significant spanwise variations were seen in the inlet-generated conical shock fronts. A downstream propagating disturbance influencing SWBLI unsteadiness was seen, and its velocity and frequency were quantified. The pressure fields of a half- and full-scale streamtraced inlet/isolator were found to be similar, with some differences attributed to viscous effects. Varying the unit Reynolds number had a limited effect on the pressure rise and SWBLI locations within the isolator. Global pressure fields, valuable for validation of computations, were measured over a region 8.5 to 12.5 isolator heights long. High-resolution measurements of the isolator pressure profiles and shock-train leading edge (STLE) position were made during the unstart process. The shape of the STLE front was measured with high spanwise resolution, and showed minimal spanwise variation in an rectangular isolator with curved corners. Time-resolved measurements show that the isolator pressure profile varies quickly in time, which presents challenges for sophisticated unstart detection methods. The unstarted streamtraced inlet flow had oscillatory or non-oscillatory modes, which were dependent on the mass flow rate of transverse air injection. Optimization of pressure transducer configuration for unstart detection was conducted. The PSP's high spatial resolution and lack of a priori decisions on where to collect pressure data were instrumental in the optimization procedure. The optimized configurations had 50.2% lower mean error and 33.5% lower maximum error in STLE location prediction on average. A relationship between the number of transducers used for STLE tracking and the error of the unstart detection method was found.双模态超燃冲压发动机（Dual-mode scramjets）是拓展高超声速飞行器（hypersonic vehicles）性能的极具潜力的技术。流线型进气道（streamtraced inlets）是高性能双模态超燃冲压发动机的极具吸引力的进气道设计方案。这类进气道会产生复杂的激波/边界层相互作用（shock-wave/boundary-layer interactions），目前相关研究在公开文献中尚未得到广泛关注，尤其是采用高分辨率测量手段的相关研究。超燃冲压发动机不起动（scramjet unstart）是指进气道质量捕获能力大幅下降的现象，通常发生在非设计工况运行过程中。该现象可在毫秒级时间尺度内发生，其支配的流动物理特性极为复杂，涉及大量激波/边界层相互作用。为指导合理的发动机设计并实现有效的流动控制以规避不起动现象，亟需对不起动过程进行更精准的预测与检测。 本研究将全局表面测量技术应用于轴对称及流线型进气道/隔离段组合结构，以更精准地表征激波/边界层相互作用与不起动现象。实验于AFOSR--圣母大学大型6马赫宁静风洞（AFOSR--Notre Dame Large Mach-6 Quiet Tunnel）中开展，来流马赫数为5.7，实验涵盖多种单位雷诺数条件。在隔离段尾部横向喷入空气射流，在冷流工况下诱导进气道不起动。本研究采用快速压力敏感涂料（pressure-sensitive paint, PSP）对隔离段内部的时变压力场进行测量，并通过温度敏感涂料测量结果对PSP数据进行温度补偿。此外，本研究采用背景纹影技术（background-oriented schlieren）对不起动过程中的进气道激波结构进行可视化观测。 本研究通过多种数据分析手段，验证了PSP在内部激波/边界层干扰动力学特性研究中的应用价值。实验将隔离段内平均激波位置的测量精度控制在0.1毫米以内。观测发现，进气道产生的锥形激波锋面存在显著的展向分布差异。研究还观测到一种沿下游传播的扰动，该扰动会影响激波/边界层干扰的非定常特性，并对其传播速度与频率进行了定量表征。研究发现，半尺度与全尺度流线型进气道/隔离段的压力场分布具有相似性，部分差异可归因于粘性效应的影响。改变单位雷诺数对隔离段内的压力升与激波/边界层干扰位置的影响有限。本研究在长度为8.5至12.5倍隔离段高度的区域内完成了全局压力场测量，这类数据对于计算模拟验证具有重要价值。 本研究在不起动过程中完成了隔离段压力分布与激波列前缘（shock-train leading edge, STLE）位置的高分辨率测量。通过高展向分辨率手段，本研究测得激波列前缘锋面的形状，并发现其在带圆角的矩形隔离段内展向变化极小。时变测量结果表明，隔离段的压力分布随时间快速变化，这为高精度的不起动检测方法带来了挑战。不起动后的流线型进气道流场存在振荡与非振荡两种模态，其模态类型取决于横向空气射流的质量流量。 本研究针对不起动检测的压力传感器配置方案进行了优化。PSP所具备的高空间分辨率特性，以及无需预先指定压力数据采集位置的优势，在该优化流程中发挥了关键作用。经优化后的配置方案，在激波列前缘位置预测中的平均误差降低了50.2%，最大误差平均降低了33.5%。本研究还发现，用于激波列前缘追踪的传感器数量与不起动检测方法的误差之间存在明确关联。