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

An experimental study of supersonic inlet buzz is presented. This study was carried out on a mixed compression rectangular inlet model; tests were done at Mach numbers ranging from 1.8 to 3, with and without bleed. Inlet flows were analyzed thanks to Schlieren videos and signal processing of unsteady pressure recordings. Two kinds of buzz were observed: "little buzz," which is thought to correspond to an acoustic resonance phenomenon excited by the presence of a shear layer under the cowl lip (Ferri criterion), and "big buzz," which seems to be triggered by a boundary layer separation on the compression ramps (Dailey criterion). In some cases, little buzz could be virtually suppressed by the introduction of a bleed. The frequency of big buzz is shown to be already present in the flow before the onset of large-amplitude oscillations, which suggests that the underlying mechanism of big buzz, probably linked to acoustics, already exists before buzz starts.

Experimental Study of Supersonic Inlet Buzz

Pseudo-shock waves and their interactions in high-speed intakes

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

An experimental and numerical investigation of the unsteady phenomena induced in a hydrogen-fuelled scramjet combustor under high-equivalence-ratio conditions is carried out, focusing on the processes leading up to unstart. The configuration for the study is the fuelled flow path of the HyShot II flight experiment. Experiments are performed in the HEG reflected-shock wind tunnel, and results are compared with those obtained from unsteady numerical simulations. High-speed schlieren and OH∗ chemiluminescence visualization, together with time-resolved surface pressure measurements, allow links to be drawn between the experimentally observed flow and combustion features. The transient flow structures signalling the onset of unstart are observed to take the form of an upstream-propagating shock train. Both the speed of propagation and the downstream location at which the shock train originates depend strongly on the equivalence ratio. The physical nature of the incipient shock system, however, appears to be similar for different equivalence ratios. Both experiments and computations indicate that the primary mechanism responsible for the transient behaviour is thermal choking, though localized boundary-layer separation is observed to accompany the shock system as it moves upstream. In the numerical simulations, the global choking behaviour is dictated by the limited region of maximum heat release around the shear layer between the injected hydrogen and the incoming air flow. This leads to the idea of ‘local’ thermal choking and results in a lower choking limit than is predicted by a simple integral analysis. Such localized choking makes it possible for new quasi-steady flow topologies to arise, and these are observed in both experiments and simulation. Finally, a quasi-unsteady one-dimensional analytical model is proposed to explain elements of the shock-propagation behaviour.

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Transient fluid-combustion phenomena in a model scramjet

Oscillatory Flows of Rectangular Hypersonic Inlet Unstart Caused by Downstream Mass-Flow Choking

Wind-tunnel tests were conducted to better understand the flow pattern of a hypersonic inlet at an unstable-unstarted condition that is caused by the downstream mass flow choking. High-speed shadowgraph flow visualization and instantaneous pressure measurements are employed to reveal the oscillations of the shock system and the unsteady process of the duct flow. Results show that the shock system both inside and in front of the inlet duct oscillates intermittently when the inlet is unstarted. With the increase of the downstream throttling ratio, the fundamental frequency of the unsteady flow process goes up and the shock-on-lip time goes down, but the process whereby the shock system is disgorged and swallowed remains unchanged, lasting for 15 ms. When the throttling ratio is 89 and 91%, the fundamental frequency is 8 and 23 Hz, respectively. At the smaller throttling ratio, the aerodynamic oscillatory phenomenon of the inlet can be divided into three stages: namely, mass filling up, shock system disgorging and swallowing, and near-throat flow pattern establishing and backpressure propagating. At the larger throttling ratio, the first stage disappears and the third one shortens.

Experimental Study of the Unstable-Unstarted Condition of a Hypersonic Inlet at Mach 6

Separation due to the shock-wave/boundary-layer interaction in a supersonic/hypersonic inlet always brings negative effects on its performance. A new control method of the shock-wave/boundary-layer interaction based on a two-dimensional bump is brought forward and investigated by both experimental and computational methods. First, the uncontrolled case is studied. The flow turning angle of the incident shock is 12 deg, and the freestream Mach number is 3.5. The results show that the shock-wave/boundary-layer interaction generates complex three-dimensional flow structures with significant swirling nature, which will substantially deteriorate the performance of a supersonic/hypersonic inlet. Then, the controlled case demonstrates that the shock-induced flow separation can be effectively reduced by the coupling of the precompression effect and the acceleration effect of the bump with a height of 0.33 times of the boundary-layer thickness. In addition, the efficiency of the control method for different shock-...

Control of Incident Shock/Boundary-Layer Interaction by a Two-Dimensional Bump

Wind-tunnel tests are conducted to study the characteristics of the shock train in the curved variable-section diffuser for hypersonic inlets. The test model is equipped with a forebody, a contracting entrance, a dump mixing duct, and an aft plug. Tests are performed at nominal freestream Mach numbers of 4.0, 5.0, and 6.0, and the corresponding inlet Mach numbers of the diffuser are 2.05, 2.59, and 3.06, respectively. Results indicate that at an inlet Mach number of 2.05, the surface pressure distributions in the shock train are similar at different backpressure ratios and can be well predicted by the modified Waltrup formula, but the length of the shock train is increased by 32 % due to the curved duct, the incident shock waves, and the incident expansion waves at the inlet plane. At higher inlet Mach numbers, the similarity disappears and the distributions of the surface pressure are not easy to predict. At different measurement points in the shock train, the instantaneous surface pressures vary obviously and almost synchronously, suggesting the oscillatory motions of the shock train. At an inlet Mach number of 2.59, the base frequency of the oscillations is about 19 Hz and the maximum fluctuating range of the instantaneous surface pressure is as high as 21.6% of the ideal pressure rise of the shock train. The static pressure fluctuations at different points in the shock train correlate strongly but are almost unrelated with those of the exit survey point, which indicates that the oscillation of the shock train is not likely induced by the pressure fluctuations far downstream of the shock train.

Preliminary Study of Shock Train in a Curved Variable-Section Diffuser

The cowl shock/boundary layer interaction is of great importance to the performance and operability of hypersonic inlets. However, it is usually affected by the expansion waves originated from the convex corner of the ramp surface (often called the shoulder), making the commonly used separation prediction relation unreliable. Therefore, the cowl shock/boundary layer interaction under the interference of the expansion waves is investigated by both theoretical analysis method and computational method in this paper. The results show that the expansion waves bring significant influence on the cowl shock/boundary layer interaction and four kinds of interaction processes exist when their relative position changes. When different interaction processes dominates the flowfield, the expansion waves may bring positive or negative effects on the shock/boundary layer interaction. In particular, while the cowl shock impinges near the shoulder, the interaction processes of shock–shock–expansion wave and shock–expansion ...

Hui-jun Tan

Papers

Oscillatory Flows of Rectangular Hypersonic Inlet Unstart Caused by Downstream Mass-Flow Choking

Experimental Study of the Unstable-Unstarted Condition of a Hypersonic Inlet at Mach 6

Control of Incident Shock/Boundary-Layer Interaction by a Two-Dimensional Bump

Preliminary Study of Shock Train in a Curved Variable-Section Diffuser

Influence of Expansion Waves on Cowl Shock/Boundary Layer Interaction in Hypersonic Inlets