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）性能的极具前景的技术。流线型进气道（streamtraced inlets）是高性能双模态超燃冲压发动机的热门进气道方案。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），相关研究在公开文献中尚未得到广泛开展，尤其是采用高分辨率测量的相关研究。超燃冲压发动机不起动（scramjet unstart）是指进气道捕获质量流量大幅下降的现象，通常发生在非设计工况（off-design conditions）下的运行过程中。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（flow physics）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避免不起动现象，亟需提升对不起动现象的预测与检测能力。
本研究将全局表面测量技术应用于轴对称式与流线型进气道/隔离段（axisymmetric and streamtraced inlet/isolators），以更精准地表征激波/边界层相互作用与不起动现象。实验在空军科学研究局-圣母大学大型马赫6宁静风洞（AFOSR--Notre Dame Large Mach-6 Quiet Tunnel）中开展，来流马赫数为5.7，涵盖多种单位雷诺数（unit Reynolds numbers）工况。通过在隔离段下游段（aft portion of the isolator）进行空气横向射流喷射（Transverse jet injection），在冷流工况（cold-flow conditions）下诱导产生进气道不起动现象。本研究采用快速压敏涂层（Fast pressure-sensitive paint, PSP）测量隔离段内部的时间分辨压力场（time-resolved pressure field），并通过温敏涂层测量实现温度补偿。同时利用背景纹影技术（Background-oriented schlieren）可视化不起动过程中的进气道激波结构。
本研究通过多种数据分析方法，验证了压敏涂层用于探究内部激波/边界层相互作用动力学特性的实用性。研究将隔离段内的平均激波位置（mean shock locations）定位精度控制在0.1毫米以内。实验观察到进气道产生的锥形激波面（conical shock fronts）存在显著的展向变化（spanwise variations）。同时观测到一种影响激波/边界层相互作用非定常特性的向下传播扰动，并对其传播速度与频率进行了量化分析。研究发现，半比例与全比例流线型进气道/隔离段的压力场特征相似，部分差异可归因于粘性效应（viscous effects）。改变单位雷诺数对隔离段内的压力升幅（pressure rise）与激波/边界层相互作用位置的影响有限。
全局压力场测量覆盖了长度为8.5至12.5倍隔离段高度（isolator heights）的区域，该类测量对计算模型验证具有重要价值。本研究在不起动过程中完成了隔离段压力剖面与激波列前缘（shock-train leading edge, STLE）位置的高分辨率测量。通过高展向分辨率测量得到了激波列前缘的轮廓形态，结果显示在带圆角的矩形隔离段内，激波列前缘的展向变化极小。时间分辨测量结果表明，隔离段压力剖面随时间快速变化，这为高精度不起动检测方法带来了挑战。不起动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喷射的质量流量。本研究还开展了用于不起动检测的压力传感器配置（pressure transducer configuration）优化工作。压敏涂层的高空间分辨率特性，以及无需预先确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。经优化的传感器配置可使激波列前缘位置预测的平均误差降低50.2%，最大误差降低33.5%。本研究还揭示了用于激波列前缘追踪的传感器数量与不起动检测方法误差之间的关联关系。