Recent research progress on unstart mechanism, detection and control of hypersonic inlet

To provide information for the detection, prediction, and control of the inlet unstart, the entire process from a started status to an unstarted status of a generic two-dimensional hypersonic inlet is studied experimentally at Mach 5. The movement of a flow plug at the exit of the duct is used to gradually increase the throttling to simulate the unstart process caused by the excessive heat release in the combustor. Simultaneous high-speed schlieren imaging and dynamic surface pressure measurements are used to record the unsteady flow structures and surface pressures of the unstart process. According to the internal and external flows, the unstart process can be divided into four stages, namely, shock train in the combustor, shock train in the isolator, separation bubble in the throat, and unstart. The transient flow patterns of each stage are substantially different and the corresponding dynamics pressures also have prominent time-frequency features, which make the detection and prediction of the inlet unstart based on dynamic pressures possible. Since the sensor placed at the end of the top surface of the inlet contract part, marked by C1, obtains the most abundant time-frequency characteristics which can be used to discern the different stages of the inlet unstart process, the location where it stays is regarded as the first choice for sensor placement to construct a practical inlet-status-monitoring system.

Experimental Investigation of the Unstart Process of a Generic Hypersonic Inlet

This paper gives an overview of a research study conducted in support of the small-scale demonstration of an active flow control system for a boundary-layer-ingesting (BLI) inlet. The effectiveness of active flow control in reducing engine inlet circumferential distortion was assessed using a 2.5% scale model of a 35% boundary-layer-ingesting flush-mounted, offset, diffusing inlet. This experiment was conducted in the NASA Langley 0.3-meter Transonic Cryogenic Tunnel at flight Mach numbers with a model inlet specifically designed for this type of testing. High mass flow actuators controlled the flow through distributed control jets providing the active flow control. A vortex generator point design configuration was also tested for comparison purposes and to provide a means to examine a hybrid vortex generator and control jets configuration. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow through the duct and the actuators. The distortion and pressure recovery were determined by 40 total pressure measurements on 8 rake arms each separated by 45 degrees and were located at the aerodynamic interface plane. The test matrix was limited to a maximum free-stream Mach number of 0.85 with scaled mass flows through the inlet for that condition. The data show that the flow control jets alone can reduce circumferential distortion (DPCPavg) from 0.055 to about 0.015 using about 2.5% of inlet mass flow. The vortex generators also reduced the circumferential distortion from 0.055 to 0.010 near the inlet mass flow design point. Lower inlet mass flow settings with the vortex generator configuration produced higher distortion levels that were reduced to acceptable levels using a hybrid vortex generator/control jets configuration that required less than 1% of the inlet mass flow.

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Boundary-Layer-Ingesting Inlet Flow Control

Review of combustion stabilization for hypersonic airbreathing propulsion

Unstart phenomena induced by flow choking in scramjet inlet-isolators

The background flow field of a scramjet isolator that accommodates a shock train contains complex compression and expansion waves, referred as “background waves,” causing large streamwise and transverse parameter gradients upstream of the shock train. Therefore, the available results of shock train research obtained by direct-connect methods might be not applicable for real scramjet isolators. Special tests are therefore performed for an inlet/isolator model. Close coupling is found between the shock train and the background shocks. The pointing direction of the leading shock switches upwards and downwards repeatedly during the upstream propagation of the shock train. Three unstable stages with substantial oscillations are also observed, interlaced with four stable stages. In addition, the interference of the background shock waves increases the sustainable back-pressure ratio and decreases the length of the shock train. However, this does not mean that the background waves in the isolator should be intensified intentionally.

Behavior of shock trains in a hypersonic inlet/isolator model with complex background waves

A hypersonic inlet with side compression has been tested at a freestream Mach number of 6.0 to enrich the understandings of hypersonic inlet unstart. A flow plug is placed at the duct exit to simulate the combustion induced high pressure and to initiate the inlet unstart. High-speed schlieren imaging and time-resolved pressure measurements are used simultaneously to record the unsteady flow structures and surface pressures of the unstart process. The inlet operates in a big buzz mode with a base frequency of 30 Hz and exhibits a series of unsteady flow patterns similar to those of rectangular hypersonic inlets in a buzz cycle when the throttling ratio is 87.4%. During the upstream propagating process of the unstart shock system, the propagation velocity in the two ends of the duct is higher than that in the middle section of the duct, with a minimum value around the isolator. Once the separation bubble induced oblique shock is expelled over the cowl lip, a supersonic reverse flow with a Mach number of 1.5...

Unstart of a Hypersonic Inlet with Side Compression Caused by Downstream Choking

To study the duct curvature effects on a shock train in rectangular hypersonic inlet/isolator models, four curved isolators are designed and tested at a freestream Mach number of 4.92. Strong inter...

Behavior of Shock Train in Curved Isolators with Complex Background Waves

Control of Cowl Shock/Boundary-Layer Interaction in Hypersonic Inlets by Bump

There are complex corner vortex flows in a rectangular hypersonic inlet/isolator. The corner vortex propagates downstream and interacts with the shocks and expansion waves in the isolator repeatedly. The supersonic corner vortex in a generic hypersonic inlet/isolator model is theoretically and numerically analyzed at a freestream Mach number of 4.92. The cross-flow topology of the corner vortex flow is found to obey Zhang’s theory [“Analytical analysis of subsonic and supersonic vortex formation,” Acta Aerodyn. Sin. 13, 259–264 (1995)] strictly, except for the short process with the vortex core situated in a subsonic flow which is surrounded by a supersonic flow. In general, the evolution history of the corner vortex under the influence of the background waves in the hypersonic inlet/isolator model can be classified into two types, namely, from the adverse pressure gradient region to the favorable pressure gradient region and the reversed one. For type 1, the corner vortex is a one-celled vortex with the ...

Shu Sun

Papers

Behavior of shock trains in a hypersonic inlet/isolator model with complex background waves

Unstart of a Hypersonic Inlet with Side Compression Caused by Downstream Choking

Behavior of Shock Train in Curved Isolators with Complex Background Waves

Control of Cowl Shock/Boundary-Layer Interaction in Hypersonic Inlets by Bump

Evolution of supersonic corner vortex in a hypersonic inlet/isolator model