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A rectangular hypersonic inlet is tested at the low-speed mode (Ma 5), the design mode (Ma 6), and the overspeed mode (Ma 7) to enrich the understanding of hypersonic-inlet unstart. A flow plug is placed at the duct exit to simulate the combustion-induced high pressure and 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 results indicate that the unstart processes are basically the same for the low-speed mode and the design mode. An unstart phenomenon dominated by the ramp-side flow separation, which exhibits violent fluctuations along the entire duct with a low frequency, is observed at these two operating modes. While operating at the overspeed mode, the strong shear layer, induced by the Mach reflection of the coalesced external oblique shock, divides the internal flow into two parts. The cowl side is filled by fully subsonic flow after the compression of the detached normal shock, and the ramp side mainly contains supersonic flow downstream of the shear-layer-induced oblique shock. Therefore, a real-time adjustment of the detached cowl shock to the downstream high pressure is formed, keeping the entire flowfield of the inlet from oscillation in a much wider range of throttling ratio. With a further increase of the throttling ratio, an oscillation dominated by the cowl-side reversed flow occurs. The oscillation phenomenon exhibits relatively gentle fluctuations, and the characteristic frequencies are mainly determined by the shock train in front of the throat section. [ABSTRACT FROM AUTHOR]

A rectangular external-compression supersonic inlet with a shock-on-lip Mach number of 2.0 is carefully studied at a freestream Mach number of 2.5. In the throttling process before the onset of buzz, the interaction between the ramp shock and the cowl-induced bow shock appears as regular reflection, Mach reflection, and λ-shaped pattern successively. It is interestingly found that the currently observed Mach reflection, which should have been broken down shortly after the emergence as reported previously, keeps fairly stable until the throttling ratio changes. In the remaining buzz stage, three types of instabilities of similar frequencies but of different oscillating behaviors and amplitudes are observed. The term medium buzz is adopted to describe the special buzz phenomenon with moderate oscillations. It is shown that both the little buzz and big buzz are characterized by the separation oscillation, whereas the medium buzz features the oscillation of the cowl-induced bow shock. Further analysis indicates that all three buzz phenomena probably share a common origin (i.e., the ramp-side flow separation). Also, it is demonstrated that the transition of buzz flows in the throttling process is attributed to the changes of disturbance strength and dominant role between two feedback loops. [ABSTRACT FROM AUTHOR]

A rectangular external-compression supersonic inlet with partially isentropic compression is experimentally studied at a freestream Mach number of 2.0. During the test, the inlet operates at the little buzz state after going through the supercritical state and the steady subcritical state. A striking feature of the little buzz state is that both the most upstream and downstream positions of the unsteady separation point of the forebody boundary layer take on a linear variation trend with the increase of throttling ratio, yielding an intensifying flow oscillation. Then, the mixed buzz phenomenon that the little and big buzz coexist appears. As the throttling ratio further rises, the occurrence probability of little buzz is reduced, and finally the big buzz state takes place. In this state, two unique transient flow patterns with the inflow and supersonic backflow coexisting at the duct entrance can be observed. Different from the prevailing viewpoint, the present little buzz belongs to the Dailey instability and shares a common origin with the big buzz in essence. Furthermore, a theoretical analysis based on Fisher's criterion indicates that a single shear layer from a λ-shaped terminal shock with an upstream Mach number below 1.60 is unable to trigger the Ferri instability of a supersonic inlet. [ABSTRACT FROM AUTHOR]

To greatly improve the supersonic inlet stability at low cost of structural weight and complexity, a novel buzz suppression strategy based on fixed-geometry air bleed is developed. It is designed to have plenty of narrow flush slots that are widely distributed along the compression surface. Using the natural pressure gradient varying with the terminal shock position, it is capable of creating self-adaptive bypass flow removal upstream of the internal duct. A strong stabilizing effect can be thus automatically produced on the subcritical flowfield by eliminating shock-induced separation and discharging excessively captured airflow. Simultaneously, the undesired air leakage at the critical regime can be naturally restricted to prevent a prohibitive performance penalty during normal operation. To verify the effectiveness, an external-compression inlet model is specially designed and carefully tested at freestream Mach numbers of 2.0 and 2.5 with an almost full exit throttle range considered (0-99.1%). Results indicate that the subcritical stable-flow range is remarkably extended from a throttle threshold of 53.7% to that of 86.4% and 73.7%, respectively, after the usage of the proposed bleed method. Moreover, intense flow instability is totally eliminated, even when the duct exit is almost closed. Further analysis reveals that the bleed flow rate at the near-critical state is not beyond 1% of the inlet flow rate for both freestream conditions. It actually causes no obvious loss of the inlet flow rate. Also, the following total pressure drop and drag increase are below 0.4%. Additionally, the observed unique buzz flow implies that the buzz origin is not necessarily limited to the two known sources, as opposed to the long-established understanding. [ABSTRACT FROM AUTHOR]

To provide information fi)r 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 pracical inlet-status-monitoring system. [ABSTRACT FROM AUTHOR]