Hypersonic Inlet Research Articles

Assessing the Performance of Hypersonic Inlets by Applying a Heat Source with the Throttling Effect

Utilization of a heat source to regulate the shock wave–boundary layer interaction (SWBLI) of hypersonic inlets during throttling was computationally investigated. A plug was installed at the intake isolator’s exit, which caused throttling. The location of the heat source was established by analysing the interaction of the shockwave from the compression ramp and the contact spot of the shockwave with that of the inlet cowl. Shockwave interaction inside the isolator was investigated using steady and transient cases. The present computational work was validated using previous experimental work. The flow distortion (FD) and total pressure recovery (TPR) of the inflows were also studied. We found that varying the size and power of the heat source influenced the shockwaves that originated around it and affected the SWBLI within the isolator. This influenced most of the performance measures. As a result, the TPR increased and the FD decreased when the heat source was applied. Thus, the use of a heat source for flow control was found to influence the performance of hypersonic intakes.

Effects of Cowl Lip Angle on Starting Performances for a Mixed-Compression Two-Dimensional Axisymmetric Hypersonic Inlet

The third quasi-uniform B-spline was adopted to design the cowl profiles with various cowl lip angles. Starting performances were simulated by the numerical method and the results showed that the starting processes for different cowls can be classified into two types. In the first type, the inlet starts when the opened large-scale separation zone is swallowed; in the second type, the opened separation zone enters the internal contraction region to form a closed separation zone, which is then swallowed to enable the start. In addition, inlets with small and large cowl lip angles exhibit the first and second types of the starting process, respectively. Moreover, there is an optimal cowl lip angle to achieve the best starting performance. For the inlet investigated herein, the starting Mach number is the lowest when the cowl lip angle is 9°. In addition, as the internal contraction ratio increases, the optimal cowl lip angle gradually decreases and the starting performance deteriorates.

Design method for hypersonic bump inlet based on transverse pressure gradient

Design method for hypersonic bump inlet based on transverse pressure gradient

Control of shock-wave/boundary-layer interaction using a backward-facing step

Flow separation induced by the shock-wave/boundary-layer interaction is detrimental to the performance of a hypersonic inlet. This paper develops a new method to control the shock-wave/boundary-layer interaction using a backward-facing step. The backward-facing step is placed on the shock-wave generator with a deflection angle of 14 deg. The shock-wave generator and a flat plate construct a simple hypersonic inlet. The control ability and mechanism are numerically studied using Reynolds Averaged Navier–Stokes equations at a freestream Mach number of 5. The results suggest that the backward-facing step can effectively suppress the flow separation. The reason is that the backward-facing step replaces the strong shock-wave in the baseline inlet with two weak ones. Consequently, the separation bubble in the controlled inlet is much smaller than that in the baseline inlet. Moreover, a parameter study is performed to analyze the effect of the design parameters of the backward-facing step on the control ability. The wave systems in the controlled inlet are classified into three types based on the results. However, only the effective wave system can suppress the shock-wave/boundary-layer interaction.

Experimental investigations of hypersonic inlet unstart/restart process and hysteresis phenomenon caused by angle of attack

In this study, the unstart/restart process of a hypersonic inlet caused by the angle of attack (AOA) is experimentally performed at a Mach number of 5.0, and the unsteady flow characteristics during the test are investigated. The flow characteristics are simultaneously recorded using high-frequency pressure measurement and high-speed schlieren imaging. The results show that the inlet enters the unstarting state at AOA = 8.4° but restarts at AOA = 0.3°, exhibiting an obvious hysteresis phenomenon. During the unstart process, the forebody shock gradually approaches the body and then impinges on the inside of the cowl with the increasing angle of attack, indicating that the inlet works under an overrated mode. As the angle of attack increases further, the airflow chokes, and a high-pressure zone is first formed in the exit of the duct. Then, the high-pressure signal propagates upstream in the duct, leading to the emergence of an inlet unstart. When the inlet enters the unstarted pattern, a typical local unstarted flow field is formed. Low-amplitude oscillations occur in the internal contraction region, but further propagation to the isolator is weak. During the restart process, when the angle of attack decreases from 3.2° to 0.3°, two kinds of big buzz patterns, namely, intermittent high-amplitude oscillation and periodic high-amplitude oscillation, appear successively. The intermittent high-amplitude oscillation is characterized by the intermittent occurrence of a high-amplitude oscillation after the low-amplitude oscillations for a random length of time. In the periodic high-amplitude oscillation, the separation bubble is repeatedly disgorged and swallowed, resulting in a dominant fluctuation frequency in the entire duct. Furthermore, as the angle of attack decreases, the dominant frequency gradually rises from 280 Hz to 339 Hz. This work provides new insights into the flow evolutions and unsteady behaviors of hypersonic inlet unstart/restart owing to the AOA and provides potential applications on the detection and prediction of hypersonic inlet unstart/restart.

Flow Characteristics of a Mixed Compression Hypersonic Intake

The flow field in a two-dimensional hypersonic mixed-compression inlet in a freestream Mach numbers of M∞ =2.0, 3.0, and 5.0 are numerically solved to understand the effect of throat area variation. The exit area ratio variation is simulated by placing a plug insert at different axial locations at the exit of the model. The flow field is achieved computationally by solving the Reynolds Averaged Navier-Stokes equations in a finite volume framework. For each flow condition, the variation in shock structure is analyzed and the variation of the oblique shock wave angle with the mass flow rate is calculated theoretically and compared with the present CFD analysis. The variation in oblique shock angle is calculated in terms of the mass flow rate by considering the capture area and spillage flow through the inlet. The theoretical results suggest that the method can predict the inlet operating conditions at different freestream Mach numbers and area ratios. This method can quantify the reduction in mass flow rate due to the throttling effect by analyzing the flow field shock pattern. The effects of various important performance parameters such as free stream Mach number, total pressure recovery, and mass flow ratio were then numerically investigated. As the Mach number is increased, the total pressure recovery is reduced, but the maximum value of the mass flow rate is increased. The analysis is also focused on the effect of throat area variation on performance parameters at each Mach number. The characteristic curve of the inlet is then obtained for each free stream Mach number.

Multi-Objective RANS Aerodynamic Optimization of a Hypersonic Intake Ramp at Mach 5

The work describes a systematic optimization strategy for designing hypersonic inlet intakes. A Reynolds-averaged Navier-Stokes database is mined using genetic algorithms to develop ideal designs for a priori defined targets. An intake geometry from the literature is adopted as a baseline. Thus, a steady-state numerical assessment is validated and the computational grid is tuned under nominal operating conditions. Following validation tasks, the model is used for multi-objective optimization. The latter aims at minimizing the drag coefficient while boosting the static and total pressure ratios, respectively. The Pareto optimal solutions are analyzed, emphasizing the flow patterns that result in the improvements. Although the approach is applied to a specific setup, the method is entirely general, offering a valuable flowchart for designing super/hypersonic inlets. Notably, because high-quality computational fluid dynamics strategies drive the innovation process, the latter accounts for the complex dynamics of such devices from the early design stages, including shock-wave/boundary-layer interactions and recirculating flow portions in the geometrical shaping.

Numerical Study on the Non-Oscillatory Unstarted Flow in a Scramjet Inlet-Isolator Model

For successful scramjet engine operations, it is important to understand the mechanism of the inlet unstart phenomenon. Among various unstarted flow patterns in hypersonic inlets, the mechanism of low-amplitude oscillatory unstarted flow is still unclear. Therefore, in the present study, the flow characteristics of non-oscillatory unstarted flow in a scramjet inlet-isolator model are studied by using numerical analysis with the RANS-based OpenFOAM solver. In the numerical results, the amplitude of pressure oscillation and the average pressure near the model outlet are in good agreement with experimental results. In the detailed analysis of the results, it is found that the incoming flow within the boundary layers repeatedly changes direction due to the flow blockage at the end of the model. In these direction-changing processes, recirculation zones near the walls irregularly influence the choked flow zones at the rear part of the model. These irregular behaviors result in non-oscillatory unstarted flow. Additionally, the main differences between the high-amplitude oscillatory unstarted flow and non-oscillatory unstarted flow are addressed.

Numerical investigation of the influence of aerothermoelastic dynamic response on the performance of the three-dimensional hypersonic inlet

The influence of aerothermoelastic dynamic response on the performance of the three-dimensional hypersonic inlet is investigated in this paper. A dynamic aerothermoelastic analysis framework is developed. The reliability of the framework is verified. The effects of aerodynamic heating on the dynamic response, the effects of the aerothermoelastic dynamic response of the inlet on the flow field structure, and the performance parameters are studied. The results indicate that the static deformation of the leading edge structure is large while the vibration amplitude is small. Meanwhile, the vibration amplitude of the trailing edge structure is larger. The aerothermoelastic dynamic response changes the shock wave structure near the exit, strengthens the shock wave intensity, increases the length of the separated region, and changes the flow field of the exit. Simultaneously, the flow field structure also experiences obvious dynamic changes at different moments. The dynamic response increases the time-averaged flow coefficient, reduces the time-averaged total pressure recovery coefficient, and improves the time-averaged reverse pressure ratio. At the same time, the dynamic response leads to the fluctuation of performance parameters, especially the large fluctuation range of the reverse pressure ratio at the exit.

CFD investigation of shock boundary layer interaction in hypersonic flow and flow control using micro ramps

PurposeThis study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.Design/methodology/approachTwo geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.FindingsThe array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.Originality/valueThis paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

Hypersonic pressure-controllable bump based on an improved permeable-boundary method

This paper presents a novel design method of pressure-controllable bump for hypersonic aircraft forebody. The new developed method can effectively resolve the tradeoff among boundary layer diversion, flow uniformity, and external drag reduction. The classical permeable-boundary method is improved by coupling with the radius-based function, through which the prescribed surface pressure distribution can inversely generate the bump. The improved permeable-boundary method is available for 3D unstructured mesh which increases the solution efficiency and robustness. The accuracy of the method is evaluated through the comparison with experimental results. In addition, five principles to arrange the pressure distribution are proposed. Then, a new pressure-controllable bump is designed. Compared with the typical streamline-tracing bump, the new bump is 44.9% lower in height while diverting identical low kinetic energy flow. According to the space occupied with high kinetic energy flow, the uniform region to preset the inlet of the new bump is 58.0% wider than that of the typical bump. Moreover, the usage efficiency of high kinetic energy flow is 1.5% higher. The new bump is less convex and its equivalent lift-to-drag ratio is two times greater than the typical bump. This research confirms that the new pressure-controllable bump shows better integrating capacity with the hypersonic inlets than the typical bump. The current design method is based on an inviscid condition; the method considering the viscous effects will be further studied in the future.

Prediction of aerothermal characteristics of a generic hypersonic inlet flow

The accurate prediction of aerothermal surface loading is of paramount importance for the design of high speed flight vehicles. In this work, we consider the numerical solution of hypersonic flow over a double-finned geometry, representative of the inlet of an air-breathing flight vehicle, characterized by three-dimensional intersecting shock-wave/turbulent boundary-layer interaction at Mach 8.3. High Reynolds numbers ($Re_L \approx 11.6 \times 10^6$ based on free-stream conditions) and the presence of cold walls ($T_w/T_o \approx 0.27$) leading to large near-wall temperature gradients necessitate the use of wall-modeled large-eddy simulation (WMLES) in order to make calculations computationally tractable. The comparison of the WMLES results with experimental measurements shows good agreement in the time-averaged surface heat flux and wall pressure distributions, and the WMLES predictions show reduced errors with respect to the experimental measurements than prior RANS calculations. The favorable comparisons are obtained using an LES wall model based on equilibrium boundary layer approximations despite the presence of numerous non-equilibrium conditions including three dimensionality, shock-boundary layer interactions, and flow separation. Lastly, it is also demonstrated that the use of semi-local eddy viscosity scaling (in lieu of the commonly used van Driest scaling) in the LES wall model is necessary to accurately predict the surface pressure loading and heat fluxes.

Parameterization and optimization design of a two-dimensional axisymmetric hypersonic inlet

Optimization method, as a promising way to improve inlet aerodynamic performance, has received increasing attention. The present research is undertaken to design a two-dimensional axisymmetric hypersonic inlet using parametric optimization. The inlet configuration is parameterized and optimized in consideration of total pressure recovery and starting performance. A Pareto front is obtained by solving the multi-objective optimization problem. Then, the flow structures of the optimized inlets are analyzed and the starting performances are evaluated. Results show that the total pressure loss mainly occurs in the internal contraction section, especially near the inlet entrance, and therefore the total pressure recovery coefficient can be greatly improved by decreasing external compression. As a result, the guidance for designing high-performance inlets is concluded. Besides, it is found that as the internal contraction ratio increases, the inlet starting ability becomes worse, which attributes to the larger separation bubble at the inlet entrance. Finally, the total pressure recovery coefficient and the starting Mach number of the optimized inlets are obtained, which can be a reference for engineering design.

Restart Processes of Rectangular Hypersonic Inlets with Different Internal Contraction Ratios

In this paper, the restart processes of two generic rectangular hypersonic inlets with different internal contraction ratios have been experimentally studied under a freestream Mach number of 5.0. The experimental results show that without downstream throttling, both inlets start successfully after the flowfield establishment of the wind tunnel. To obtain the restart processes of the two inlets, they are initially set at an unstarted state by a large downstream throttling ratio. Then, the throttling ratio is continuously decreased. For the inlet with a small internal contraction ratio, the restart process undergoes four stages: stable buzz, intermittent buzz, reestablishment of wave system in the inlet, and unthrottled flow; whereas for the inlet with a large contraction ratio, the internal flow cannot be normally reestablished after the intermittent buzz stage because a large-scale separation bubble consistently stands ahead of the inlet. That is, it cannot be restarted. In the stage of the intermittent buzz, there is a probability of the appearance of a quasi-steady flowfield when the pressure in the inlet reaches its maximum value, and the separation shock almost impinges on the cowl lip. For both inlets, the probability of buzz decreases continuously with time, and then it suddenly disappears. Furthermore, for the inlet with a large contraction ratio, two kinds of flowfields (that is, started or unstarted) occur at the same downstream throttling ratio and upstream flow conditions, which indicates that the inlet operates in the dual-solution region and the actual operating state of the inlet significantly depends on the historical flow. Therefore, the evaluation of hypersonic inlet starting capability must be examined carefully.

Transitions of shock interactions on V-shaped blunt leading edges

Abstract

Magnetohydrodynamic Control of Hypersonic Separation Flows

Magnetohydrodynamic (MHD) control of hypersonic laminar separation flows is investigated in this paper. A series of numerical simulations over various geometry configurations, namely, a compression corner and a double wedge ramp hypersonic inlet, have been conducted by application of an external electromagnetic field. Results show that the performance of MHD separation flow control is mainly determined by flow acceleration of the Lorentz force directed in the streamwise direction. The Joule heating term always brings negative effects on the MHD separation flow control and increased the static pressure locally, where the electromagnetic field is applied. With an external electromagnetic field applied, the low velocity fluid in the boundary layer can be accelerated. Moreover, there exists a best location for the MHD zone to be applied and completely eliminate the separation of the flow from the surface.

On the unsteady throttling dynamics and scaling analysis in a typical hypersonic inlet–isolator flow

The flow field in a two-dimensional three-ramp hypersonic mixed-compression inlet in a freestream Mach number of M∞ = 5 is numerically solved to understand the unsteady throttling dynamics. Throttling conditions are simulated by varying the exit area of the isolator in the form of plug insets. Different throttling ratios between 0 ≤ ζ ≤ 0.7 in steps of 0.1 are considered. No unsteadiness is observed for ζ ≤ 0.2, and severe unsteadiness is found for 0.3 ≤ ζ ≤ 0.7. The frequency of unsteadiness (f) increases rapidly with ζ. As ζ increases, the amount of reversed mass inside the isolator scales with the frequency and the exit mass flow rate. A general framework is attempted to scale the unsteady events based on the gathered knowledge from the numerical study. The inlet–isolator flow is modeled as an oscillating flow through a duct with known upstream design conditions such as the freestream Mach number (M∞) and the isolator inlet Mach number (Mi). Factors such as the mass occupied by the duct volume, the characteristic unsteady frequency, the throttling ratio, and the exit mass flow rate through the duct are used to form a non-dimensional parameter β, which scales with the upstream design parameter ξ = Mi/M∞. The scaling parameters are further exploited to formulate a semi-empirical relation using the existing experimental results at different throttling ratios from the open literature. The unsteady frequencies from the present two-dimensional numerical exercise are also shown to agree with the proposed scaling and the resulting semi-empirical relation.

Feasibility of employing the restarting process to evaluate the self-starting ability for hypersonic inlets

Accurate evaluation of hypersonic inlet self-starting performance is very crucial for the stable and efficient operation of scramjet engines. However, few ground hypersonic wind tunnels could provide a continuously and slowly accelerating free stream conditions to directly validate the self-starting ability of hypersonic inlets so far. Some experimental restarting approaches have been explored to evaluate the self-starting performance of hypersonic inlets. To clarify the feasibility of such approaches, the self-starting process and the restarting process caused by withdrawing plug at the exit of a two-dimensional hypersonic inlet were studied using unsteady numerical simulation in this paper. The results show that, even though the freestream Mach number (5.54) of the restarting process is lower than the self-starting Mach number (5.715), the hypersonic inlet can finally rebuild the started flow field owing to the significantly higher moving speed downstream across the inlet internal contraction section caused by the big buzz. Hence, the additional unsteady effect should be suppressed as far as possible during the initial pre-blockage process in the restarting experiment applied to evaluate the self-starting performance of hypersonic inlets.

Self-starting Simulation of a Hypersonic Inlet with Variable Free Stream Condition

Self-starting Simulation of a Hypersonic Inlet with Variable Free Stream Condition

Integration of inward-turning inlet with airframe based on dual-waverider concept

This paper proposes an inward-turning-inlet–forebody integration methodology based on the dual-waverider concept. For the design concept and the designed flight conditions, both the airframe and inlet benefit from the waverider's high lift-to-drag ratio and the ideal compression performance of its propulsion system. The critical issue is to construct a complex three-dimensional shock wave that accounts for both the inlet and airframe. A method is then proposed for connecting the axisymmetric flow fields of the internal and external flows. The compression surface of the integrated vehicle is then generated from the streamlines obtained with a series of axisymmetric basic flows. Finally, the integration configuration is verified using computational methods, and its viscous effects are analyzed under both on- and off-design conditions via numerical simulations. The results show that a configuration designed using the proposed approach has a high corresponding lift-to-drag ratio and high hypersonic inlet performance.