Mixed-compression Supersonic Inlet Research Articles

Study on coupling characteristics of supersonic inlet and pulse detonation combustor

Coupling model for two-dimensional mixed-compression supersonic inlet with Mach number of 2.5 at design point and pulse detonation combustor was established, and three working schemes were designed to investigate multi cycle working process by using unsteady numerical simulation with chemical reaction. The results show that the following two measurements can both restrain the back-propagating compression wave and broaden the stability margin of the inlet and increase the frequency, one of which is the pressure relief at the outlet of the combustor, the other is valve at the inlet of the combustor. Nevertheless, the former will lose thrust, while the latter can greatly increase thrust. Meanwhile, the valve will reflect the compression wave in the production; increase the turbulence in the combustor, and make the gas entrained into the diffuser section of the inlet at the moment of opening the valve. However, the reflected compression wave in the hot production will not affect the velocity and shape of the detonation wave. Furthermore, in the period of detonation wave initiation and propagation, the blockage ratio of physical obstructions at the inlet of the combustor should be expanded as much as possible. At the same time, the flow at the outlet of the inlet should not be blocked to increase the flow resistance.

Suppression of flow response hysteresis in the throttling/unthrottling process for supersonic inlet

The flow response hysteresis in a mixed-compression supersonic inlet's throttling/unthrottling process is unconducive to its practical engineering applications. To suppress such hysteresis, herein, a flow control method based on a distributed bleed system in the inlet's internal contraction part was proposed and applied to a rectangular supersonic inlet with a designed Mach number of 4.0 and an internal contraction ratio of 1.5. All tests were conducted in a supersonic wind tunnel at a freestream Mach number of 2.9, and the relevant test data were obtained using the high-speed schlieren and dynamic pressure measurement systems. The results indicate that the distributed bleed system improved the uncontrolled inlet's unstart throttling ratio from 43.8% to 47.3% and the restart throttling ratio from 29.4% to 47.1%, and it essentially eliminated the flow response hysteresis during the throttling/unthrottling process. Moreover, the hysteresis mechanism of the uncontrolled inlet, dominated by the separation-induced flow structures in the internal contraction part, was investigated to help elucidate the targeted control mechanism of the bleed system that suppresses the hysteresis. Finally, the specific flow phenomena in the oscillation of the controlled inlet were examined to study the bleed system's impact on the unsteady flowfield, including the coupled flow of the mainstream with the bleed system and complex wave motion structures.

Novel Simplified Numerical Simulation Method for Modeling Bleed Holes in Supersonic Inlets

A novel simplified numerical simulation method for bleed holes in supersonic inlets was developed to estimate the bleed mass flow rate and simulate the flowfield structure in the bleed region. First, the Prandtl–Meyer expansion theory was employed to calculate the location of the barrier shock and reconstruct the flow structure inside the bleed hole. Next, a novel discrimination algorithm was developed to distinguish the grids connected to the bleed hole in the bleed zone instead of directly modeling the hole opening and plenum chamber. Finally, the bleed mass flow rate was obtained based on the boundary-layer properties in the bleed region and the fitting scaled coefficients. This method was also evaluated by performing numerical simulations for a series of different bleeding situations. For all the examined cases, the simulation results were in good agreement with the experimental data, and the number of grid points’ requirement for modeling the bleed hole as compared with the fully resolved method was considerably reduced. Moreover, the proposed method was successfully applied to simulate porous bleed systems in a mixed-compression supersonic inlet. This illustrates that the proposed method is expected to be a useful engineering tool for designing the bleed system of the supersonic inlets.

Comparative study on single-incident and dual-incident shock wave/turbulent boundary layer interactions with identical total deflection angle

Interactions between the boundary layer and two successive incident shock waves (ISWs) often occur in the supersonic mixed-compression inlets. However, the flow mechanism involved in such interactions has been studied rarely. In this study, we investigate experimentally and analytically the turbulent boundary layer separation flow induced by the single ISW and dual ISWs at the identical total deflection angles in a Mach 2.73 flow. Schlieren photography, wall pressure measurement and surface oil-flow visualisation are employed to diagnose the flow field. Experiments with the impingement points of the two ISWs intersecting on the bottom wall exhibit a separated flow with a triangle-like separation bubble, namely the first kind of dual-ISW/turbulent boundary layer interaction (ISWTBLI). Comparative studies show that various flow features of this kind of dual-ISWTBLI, including the extent of the separation region, pressure distribution and surface-flow topological structures, are nearly identical to those of the single-ISWTBLI with an identical total deflection angle. As the distance between the two ISWs increases, the shape of the separation region in the dual-ISWTBLI changes from triangle-like to quadrilateral-like, and the height of the separation region decreases accordingly, forming the second kind of dual-ISWTBLI. Furthermore, an inviscid model is developed for the dual-ISWTBLI to describe the complex shock wave system and elucidate the cause of a quadrilateral-like separation bubble in the second kind of dual-ISWTBLI. Moreover, based on a previous work by Souverein et al. (J. Fluid Mech., vol. 714, 2013, pp. 505–535) on the single-ISWTBLI, a modified scaling method is established for the first kind of dual-ISWTBLI.

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.

Buzz characteristics and separation bubble dynamics in supersonic intake

The flow field in an axisymmetric supersonic mixed-compression inlet is numerically solved to understand the separation bubble dynamics at different freestream Mach numbers of M∞ = 2.0, 3.0, 5.0, and throttling ratios of ψ = 0.515, 0.40, and 0.35. The two-dimensional compressible Reynolds Averaged Navier Stokes equations were solved in a finite volume framework. The computational results were validated with the wind tunnel experimental data for steady as well as unsteady operating conditions. Throttling conditions are simulated by varying the exit area of the model in the form of plug inserts. The analysis of the flow features in the complete computation domain indicates the formation of multiple separation bubbles near the exit plug as well as near the isolator entrance of the inlet. The frequency of unsteadiness (f) associated with the ramp surface pressure decreases with a decrease in throttling ratio ψ. The possibility of a common origin that is regardless of any buzz criteria is analyzed by considering separation and reattachment zones in the flow field in terms of wall shear stress variation. Inside the supersonic intake, the multiple shock waves interact with the wall boundary layer that can lead to a large separation bubble. This bubble is always unstable inside the intake flow passage and can generate the source of instability. The propagation of a pressure disturbance is analyzed by the temporal variation of the pressure on the ramp wall. The cross-correlation coefficient Cij(t) of the pressure signal on the ramp surface suggests the upstream propagation of the disturbance generated from the separation zone near the exit plug. The numerical Schlieren provides further information on the formation and propagation of the separation bubble near the exit plug.

Study of the Influence of the Temperature Field’s Nonuniformity on the Pressure Pulsations in the Inlet Duct Using the RANS/ILES Method

The paper is devoted to the RANS/ILES high-resolution method application for studying the influence of a nonuniformity of the inflow temperature field on the throttle characteristics of a supersonic mixed compression inlet, and the intensity and spectral characteristics of pressure pulsations in the inlet duct. Five cases of inflow parameters are investigated, three of which have a nonuniform temperature field. The sound pressure level distributions in the inlet duct and the pressure pulsation spectra at points on the inlet walls are obtained.

Control of a Supersonic Inlet in Off-Design Conditions with Plasma Actuators and Bleed

Supersonic inlets are a key component of present and future air-breathing propulsion systems for high-speed flight. The inlet design is challenging because of several phenomena that must be taken under control: shock waves, boundary layer separation and unsteadiness. Furthermore, the intensity of these phenomena is strongly influenced by the working conditions and so active control systems can be particularly useful in off-design conditions. In this work, a mixed compression supersonic inlet with a double wedge ramp is considered. The flow field was numerically investigated at different values of Mach number. The simulations show that large separations appear at the higher Mach numbers on both the upper and lower walls of the duct. In order to improve the performances of the inlet two different control strategies were investigated: plasma actuators and bleed. Different locations of the plasma actuator are considered in order to also apply this technology to configurations with a diverter which prevents boundary layer ingestion. The potential of the proposed solutions is investigated in terms of total pressure recovery, flow distortion and power consumption.

Supersonic inlet buzz detection using pressure measurement on wind tunnel wall

Feasibility of an innovative buzz detection technique through measuring the static pressure outside a mixed-compression supersonic inlet is studied. The buzz is an instability phenomenon that occurs almost in all supersonic inlets. During the buzz, shock oscillation along with pressure and mass flow fluctuations affects the performance characteristics of the inlet. The main objective of this paper is to introduce a simple and easy-to-implement method for investigation of the buzz phenomenon in a supersonic inlet. The experimental data for far field-based are compared with those of the model-based one at free stream Mach numbers of 1.8, 2.0, and 2.2 and at zero degrees angle of attack for a mixed-compression inlet. The results show that this technique can measure exact value of buzz frequency as well as its onset. The present method uses pressure data obtained from the wind-tunnel wall instead of measuring the pressure inside or on the model surfaces which in most cases is very hard. However, to sense flow oscillations, caused by the buzz onset the sensors on the wind-tunnel wall must be located downstream of the point where the oblique shock from the spike impinges on the tunnel wall. The location of the measurement point as well as the distance between the sensor and the origin of the shock wave system are very important.

Numerical Investigation of Back Pressure and Free-stream Effects on a Mixed Compression Inlet Performance

Inlet Performance has an important role in the operation of air-breathing propulsion systems. In this study, performance of a supersonic axisymmetric mixed-compression inlet in the supercritical operating condition is numerically studied. The effects of free-stream Mach number and engine-face pressure on performance parameters, including mass flow ratio, drag coefficient, total pressure recovery, and flow distortion are investigated. To this sake, a multi-block density-based finite volume CFD code is developed and Reynolds-averaged Navier-Stokes equations with Spalart-Allmaras one-equation turbulence model is employed. The code is validated by comparing numerical results against other computational results and experimental data for two test cases of inviscid flow in a two-dimensional mixed-compression inlet and flow in an external compression inlet. Finally, the code is utilized for investigation of a specific supersonic mixed-compression inlet with the design Mach number of 2.0 and length to diameter ratio of 3.4. Results revealed that the increment of free-stream Mach number leads to decrease in total pressure recovery and drag coefficient, while mass flow ratio and flow distortion increase. The effects of engine-face pressure on performance parameters showed that by increasing the engine-face pressure, mass flow ratio and drag coefficient remain constant while total pressure recovery increases and flow distortion decreases.

English

English

A study of unsteady-state operating conditions of a supersonic inlet by the RANS/ILES method

Using the hybrid RANS/ILES method, a flow was investigated under unsteady-state conditions in a rectangular mixed-compression supersonic inlet at Md = 2. The computations were made for Mach numbers of the incident flow M0 = 1.8, 2, and 3. The geometry of two variants was investigated, viz., with the boundary layer bleed system and without it. Calculations were performed on the meshs containing (1.69–1.78) × 106 cells. The flow rate through the supersonic inlet was varied within a wide range. Under most of the conditions investigated, the flow in the duct was unsteady as a consequence of separation of the boundary layer upon its interaction with the shock waves. For all geometric variants investigated and M0, the throttle characteristics were constructed by the averaged flow parameters as well as the dependences of the flow rate through the boundary layer bleeding system and the static pressure pulsation level on the air inlet throttle ratio. Comparison of the computed results with the experimental data showed a good agreement in terms of both the averaged flow parameters and the pulsation characteristics.

Modeling the Effect of Stability Bleed on Back-Pressure in Mixed-Compression Supersonic Inlets

To stabilize the terminal normal shock on high-static pressure at outlet, called back-pressure pout, stability bleed slots are used in the throat of mixed-compression supersonic inlets. In this paper, a model for the functional relation between the bleed flow rate mbl and back-pressure pout is established based on a bleed flow rate model (BFRM) in order to study the effect of stability bleed on the back-pressure in mixed-compression supersonic inlets. Given the inlet flow parameters Min, pin*, and Tin*, the plenum pressure ppl at slots' outlet, the terminal normal shock position xs in this model, the bleed flow rate mbl, Mach number M¯out, and back-pressure pout were derived one by one from the basic laws of conservation. To study the effect of plenum pressure ppl on subsonic flow of the divergent section behind the terminal normal shock, a correction coefficient κ is introduced to modify the Mach number M¯out. Furthermore, numerical simulations based on Reynolds-Averaged Navier–Stokes equations were performed to analyze the functional relation between the bleed flow rate mbl and back-pressure pout. Computational fluid dynamics (CFD) results show that the present model agrees with the data.

Morphing Supersonic Inlet with Deforming Air Cell

A mixed-compression supersonic inlet with a variable throat can solve the contradiction of the starting capability at low Mach numbers and the total pressure recovery performance at high Mach numbers. However, the variation of the throat area is rather difficult for an axisymmetric supersonic inlet. A new method based on a deforming air cell to control the throat area of an axisymmetric inlet is brought forward and investigated by both experimental and computational methods. The morphing behavior of the deforming air cell in a supersonic wind tunnel is also obtained. The results indicate that the contraction ratio of the inlet can be effectively controlled by the cooperation of the air cell profile under different charging pressures and the cowl inner surface. As compared with a fixed geometry inlet, the total pressure recovery and the corresponding maximum backpressure of the morphing inlet are increased by 25.7 and 31.8%, respectively. In addition, the air cell can withstand the impact of strong oscillatory flows when the inlet experiences buzz.

A quasi one-dimensional bleed flow rate model for terminal normal shock stability in mixed compression supersonic inlet

Ninety-degree (normal) bleed slots have been used to stabilize the terminal normal shock in the throat of a mixed compression supersonic inlet. In this study, a quasi-one-dimensional bleed flow rate model, consisting of a constant-area channel with a pair of normal slots symmetrically located along the upper and lower endwalls, is developed. The bleed flow rate is shown to be a function of the terminal normal shock position within the slot. Some key factors, such as the bleed discharge coefficient, taken from the Bragg model, were derived from the basic laws of conservation for a one-dimensional simplification. Furthermore, numerical simulations based on Reynolds-averaged Navier–Stokes equations were performed to analyze the flow characteristics around the bleed slots. The predictions of the bleed flow rate model agree well with computational fluid dynamics results. This method may be helpful to predict the stability of the terminal normal shock in mixed compression supersonic inlets.

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, α, 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, α, 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 character- ristics 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.

Euler flow in a supersonic mixed-compression inlet

Numerical simulation of a two-dimensional mixed compression supersonic inlet is carried out by solving unsteady compressible Euler equations via a stabilized finite element method. The geometry of the inlet is similar to the one used by Anderson and Wong for experimental investigation for Mach 3 flow. The computations are capable of simulating the start-up problems associated with the inlet. The critical back pressure for the successful operation of the inlet is computed. The effect of inlet back pressure on the total pressure recovery and the flow distortion level is analysed. Contrary to the popular belief, it is found that in addition to the throat to inlet capture area ratio, the ramp geometry close to the throat plays an important role in the start-up dynamics. It is demonstrated via simulations that, everything else being same, the geometries of ramp upstream of the throat that are associated with a curvature higher than a certain threshold, result in unstarting the intake.

Aeroelastically deflecting flaps for shock/boundary-layer interaction control

An aeroelastic mesoflap system has been developed to improve the downstream flow properties of an oblique shock/boundary-layer interaction. The mesoflap system employs a set of small flaps over a cavity, whereby the flaps downstream of the interaction bend downward aeroelastically to bleed the flow and the upstream flaps bend upward to re-inject this same mass flow upstream. This recirculating system requires no net mass bleed and therefore has advantages for boundary layer control in external or mixed-compression supersonic aircraft inlets. In addition, the system may be applicable in other aerospace applications where boundary-layer control can help remedy the adverse effects of shock interactions. Several mesoflap systems have been fabricated and examined experimentally to investigate their aerodynamic and structural performance. Each mesoflap is rigidly attached to a spar on its upstream end while the remainder of the flap is free to deflect aeroelastically. The flap length is nominally a few boundary-layer thicknesses in dimension, while the flap thickness is small enough to allow tip deflections that are of the order of the boundary-layer momentum thickness. Experiments were conducted for a Mach 2.41 impinging oblique shock wave interaction with a turbulent boundary layer. Spanwise-centered laser Doppler velocimeter measurements indicate that certain mesoflap designs can show significant flow improvement as compared to the solid-wall case, including increased stagnation pressure recovery and a 7% reduction in boundary layer thickness and sonic thickness. However, one drawback of the mesoflap system is the potential for fatigue, which in some cases led to microcracking followed by flap failure. Structural design improvements to alleviate and avoid this problem included a lower profile spar design, substitution of Nitinol for aluminum as the flap material, and use of stress-relieving holes at the ends of the flap cut-outs.

Modeling the Response from a Cascade to an Upstream Acoustic Disturbance

Time-accurate Euler flow field simulations for the flow through a two-dimensional cascade subjected to an upstream acoustic disturbance were used as the basis for a small disturbance model to predict the reflected response upstream of the cascade. The small disturbance model results in a linear system of algebraic equations for the properties of the reflected and transmitted disturbances. The model predicts the reflected and transmitted responses as a function of the cascade blade geometry, the disturbance strength, and the initial flow properties prior to the upstream disturbance. The predicted results from the small disturbance model were then compared with the Euler analysis results for a two-dimensional cascade. Agreement between the model and the Euler data indicated that the model was potentially useful as a basis for an outflow boundary condition for time-accurate Euler/Navier-Stokes (ENS) simulations of supersonic mixed compression inlet flows needed to determine the stability margin of an inlet that encounters an atmospheric disturbance. This boundary condition must provide an approximation of the response from the compressor by the inlet flow at the face of the compressor when a disturbance from upstream passes through the inlet and into the compressor. A new characteristic boundary condition based on the small disturbance response model was formulated and demonstrated independently in two one-dimensional Euler codes. The one-dimensional Euler codes with the new boundary condition and with existing boundary condition formulations were used to predict the reflection response for an axial compressor experiment. The new boundary condition was found to provide a significant improvement in accuracy for the reflection response of an acoustic disturbance from a compressor relative to existing outflow boundary condition models. For a supersonic mixed compression inlet, a one-dimensional Euler code was also used to demonstrate the dependence of the inlet normal-shock response and unstart tolerance on the outflow boundary condition.

Numerical investigation of supersonic inlet with realistic bleed andbypass systems

A supersonic mixed-compression inlet was designed for Mach 2.5 and aerodynamically analyzed using a GFD approach. Its ramp bleed and throat bypass systems were simulated in order to examine the physical nature of flow in the system. From the calculation, two types of unstart were predicted. One is a shock/boundarylayer interaction induced unstart. The other is usual unstart due to failure of pressure balance. A ramp bleed was found to play an important role for avoiding the former unstart. For avoiding or delaying the latter unstart, a throat bypass system plays an important role. Shock-capturing and stabilizing effects by the throat bypass were also confirmed numerically. Furthermore, the mechanism of improving pressure recovery in a shockcapturing mode, a critical mode, was examined in detail and the important role of a throat slot was revealed numerically for the first time. A streamline passing over the slot automatically constructs an optimum equivalent wail boundary in accordance with the flow conditions around the system. On the high-back-pressure-shockcapturing mode, the equivalent wall boundary reduces mainstream throat area, because of high pressure in the throat plenum chamber, and weakens the terminal shock. As a result, the total pressure recovery improves. This noteworthy result could be obtained owing to the great advantage of the CFD approach.