Experimental Investigation of a Fixed-geometry Two-dimensional Mixed-compression Supersonic Inlet with Sweep-forward High- light and Bleed Slot in an Inverted ＂X＂-type Layout

A fixed-geometry two-dimensional mixed-compression supersonic inlet with sweep-forward high-light and bleed slot in an inverted ＂X＂-form layout was tested in a wind tunnel. Results indicate： （1） with increases of the free stream Mach number, the total pressure recovery decreases, while the mass flow ratio increases to the maximum at the design point and then decreases; （2） when the angle of attack, a, is less than 6°, the total pressure recovery of both side inlets tends to decrease, but, on the lee side inlet, its values are higher than those on the windward side inlet, and the mass flow ratio on lee side inlet increases first and then falls, while on the windward side it keeps declining slowly with the sum of mass flow on both sides remaining almost constant; （3） with the attack angle, a, rising from 6° to 9°, both total pressure recovery and mass flow ratio on the lee side inlet fall quickly, but on the windward side inlet can be observed decreases in the total pressure recovery and increases in the mass flow ratio; （4） by comparing the velocity and back pressure characterristics of the inlet with a bleed slot to those of the inlet without, it stands to reason that the existence of a bleed slot has not only widened the steady working range of inlet, but also made an enormous improvement in its performance at high Mach numbers. Besides, this paper also presents an example to show how this type of inlet is designed.

Experimental investigation on unstart-restart hysteresis of a supersonic inlet during throat regulation

The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart.Hence,detailed experimental studies of such a hysteresis and its control are necessary.A throat variable supersonic inlet was designed at a shock-on-lip Mach number of 4.0 and an Internal Contraction Ratio(ICR)ranging over 1.21–2.94.Meanwhile,a distributed bleed system was proposed to suppress the hysteresis.The wind tunnel tests were conducted at Mach number 2.9.The throat regulation processes were recorded using a high-speed schlieren and dynamic pressure acquisition system.The results indicate that the unstart and restart ICRs during the uncontrolled inlet’s throat regulation process were 1.95 and 1.48,respectively,demonstrating an unstart-restart hysteresis.Four typical flowfields were summarized during the uncontrolled inlet’s restart process.The proposed bleed control increased the unstart and restart ICRs to 2.06 and 1.75,respectively,and the inlet realized the designed state as the ICR was further decreased to 1.67.The controlled inlet’s hysteresis loop was decreased compared to the uncontrolled inlet.Finally,the mechanism of the hysteresis,dominated by the entrance separation-induced wave system,was clarified.The mechanisms of the bleed control to broaden the unstart and restart boundaries and suppress the hysteresis were elucidated.

Flow response hysteresis of throat regulation process of a two-dimensional mixed-compression supersonic inlet

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock,so as to reduce the flow loss.However,excessive Internal Contraction Ratio(ICR)exposes the inlet to a greater risk of unstart,which inevitably results in a process of increasing the throat area to aid the inlet restart.In the above throat regulation process,the inlet undergoes the start,unstart,and restart states in turn.In order to reveal the flow structure and mechanism of this process,a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed.The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is+6°.Analysis was focused on the state with a freestream Mach number of 3.0.The results clearly show that the flow response hysteresis appears,and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure.In addition,three typical flow fields were analyzed,and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis.Finally,unstart and restart boundaries of the inlet were discussed,and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed.The newly predicted unstart and restart boundaries are much closer to the CFD results.

Minimization of Classification Samples for Supercritical and Subcritical Patterns of Supersonic Inlet

In order to investigate sample minimization for classification of supercritical and subcritical patterns in supersonic inlet, three optimization methods, namely, opposite one towards nearest method, closest one towards the byper-plane method and random selection method, are proposed for investigation on minimization of classification samples for supercritical and subcritical patterns of supersonic inlet. The study has been carried out to analyze wind tunnel test data and to compare the classification accuracy based on those three methods with or without priori knowledge. Those three methods are different from each other by different selecting methods for samples. The results show that one of the optimization methods needs the minimization samples to get the highest classification accuracy without priori knowledge. Meanwhile, the number of minimization samples needed to get highest classification accuracy can be further reduced by introducing priori knowledge. Furthermore, it demonstrates that the best optimization method has been found by comparing all cases studied with or without introducing priori knowledge. This method can be applied to reduce the number of wind tunnel tests to obtain the inlet performance and to identify the supercritical/subcritical modes for supersonic inlet.

Axisymmetric and three-dimensional flow simulation of a mixed compression supersonic air inlet

The flow through an axisymmetric supersonic mixed-compression air inlet has been simulated numerically to investigate the effects and the necessity of the three-dimensional(3D)modeling in comparison with the axisymmetric one.For this purpose,a supersonic inlet has been simulated numerically via axisymmetric and 3D CFD solvers,using the steady state Reynolds-averaged Navier-Stokes equations along with the SST k-ωturbulence model,for a free-stream Mach number of 2.0 and at zero degrees angle of attack.The grid for the 3D cases was a 14.4-degree sector,instead of a 360-degree domain one,with rotational periodic boundary condition for the side boundaries.The results show that both static and total pressure distributions match well with the experimental data for both the axisymmetric and the 3D simulations.If the prediction of performance parameters is the main goal of simulations,it seems that the axisymmetric simulation provides adequate accuracy,and the 3D simulation one is not the best choice.The 3D numerical simulation results in an in-depth study on the supersonic inlets,including the shock wave-boundary layer interaction,the location of the terminal normal shock wave,and consequently the separation point.For an axisymmetric supersonic inlet in an axisymmetric flow condition,3D effects are not strong enough to have a significant influence on the inlet performance for all operational conditions.However,it seems that 3D effects play an important role in both critical and supercritical operating conditions during the steady state operation.

A theoretical and 1-D numerical investigation on a valve/valveless air-breathing pulse detonation engine

The important operating characteristics of pulsed Pressure Gain Combustion(PGC)propulsion are the pressure gain of the combustor component and the propulsive performance gain of the engine.A ramjet-type valve/valveless air-breathing pulsed detonation engine with a supersonic internal compression inlet is investigated.Based on an ideal thermal cycle,the ideal equivalent pressure ratios(pcb)of the Pulsed Detonation Combustor(PDC)are obtained theoretically which are directly related with the propulsive performance of the engine.By introducing an orifice loss model into the cycles,the critical pressure drop ratios through the orifice for the PDC achieving pressure gain and the engine achieving thrust gain are studied.More influencing factors are investigated by the use of a one-dimensional(1-D)numerical simulation model.The operating characteristics of the pulse detonation engine are investigated with changes of the valve type,the inlet/outlet area ratio of the PDC,the nozzle area ratio,and flight conditions.All these factors can affect pcbof the PDC,and pcbcan be optimized by changing the geometry of the engine.The most important influence parameter is the valve type.When using an orifice-type aerodynamic valve,simulation results show that the PDC cannot achieve the pressure gain characteristics.When a supersonic internal compression inlet is introduced to the engine,whether the Pulse Detonation Engine(PDE)can achieve thrust gain comparable with that of an ideal Brayton cycle engine not only is related to the pressure gain of the combustor,but also needs to optimize the engine structure to reduce the total pressure loss.