HEATING CHARACTERISTICS OF BLUNT SWEPT FIN-INDUCED SHOCK WAVE TURBULENT BOUNDARY LAYER INTERACTION

An experimental study was conducted on shock wave turbulent boundary layer interactions caused by a blunt swept fin-plate configuration at Mach numbers of 5.0, 7.8, 9.9 for a Reynolds number range of (1.0.similar to 4.7) x 10(7)/m. Detailed heat transfer and pressure distributions were measured at fin deflection angles of up to 30 degrees for a sweepback angle of 67.6 degrees. Surface oil flow patterns and liquid crystal thermograms as well as schlieren pictures of fin shock shape were taken. The study shows that the flow was separated at deflection of 10 degrees and secondary separation were detected at deflection of theta greater than or equal to 20 degrees. The heat transfer and pressure distributions on flat plate showed an extensive plateau region followed by a distinct dip and local peak close to the fin foot. Measurements of the plateau pressure and heat transfer were in good agreement with existing prediction methods, but pressure and heating peak measurements at M greater than or equal to 6 were significantly lower than predicted by the simple prediction techniques at lower Mach numbers.

Numerical evaluation of passive control of shock wave/boundary layer interaction on NACA0012 airfoil using jagged wall

Shock formation due to flow compressibility and its interaction with boundary layers has adverse effects on aerodynamic characteristics, such as drag increase and flow separation. The objective of this paper is to appraise the practicability of weakening shock waves and, hence, reducing the wave drag in transonic flight regime using a two-dimensional jagged wall and thereby to gain an appropriate jagged wall shape for future empirical study. Different shapes of the jagged wall, including rectangular, circular, and triangular shapes, were employed. The numerical method was validated by experimental and numerical studies involving transonic flow over the NACA0012 airfoil, and the results presented here closely match previous experimental and numerical results. The impact of parameters, including shape and the length-to-spacing ratio of a jagged wall, was studied on aerodynamic forces and flow field. The results revealed that applying a jagged wall method on the upper surface of an airfoil changes the shock structure significantly and disintegrates it, which in turn leads to a decrease in wave drag. It was also found that the maximum drag coefficient decrease of around 17 % occurs with a triangular shape, while the maximum increase in aerodynamic efficiency（lift-to-drag ratio）of around 10 % happens with a rectangular shape at an angle of attack of 2.26？.

Hypersonic Shock Wave/Boundary Layer Interactions by a Third-Order Optimized Symmetric WENO Scheme

A novel third-order optimized symmetric weighted essentially non-oscillatory(WENO-OS3)scheme is used to simulate the hypersonic shock wave/boundary layer interactions.Firstly,the scheme is presented with the achievement of low dissipation in smooth region and robust shock-capturing capabilities in discontinuities.The Maxwell slip boundary conditions are employed to consider the rarefied effect near the surface.Secondly,several validating tests are given to show the good resolution of the WENO-OS3 scheme and the feasibility of the Maxwell slip boundary conditions.Finally,hypersonic flows around the hollow cylinder truncated flare(HCTF)and the25°/55°sharp double cone are studied.Discussions are made on the characteristics of the hypersonic shock wave/boundary layer interactions with and without the consideration of the slip effect.The results indicate that the scheme has a good capability in predicting heat transfer with a high resolution for describing fluid structures.With the slip boundary conditions,the separation region at the corner is smaller and the prediction is more accurate than that with no-slip boundary conditions.

STUDY OF SWEPT SHOCK WAVE AND BOUNDARY LAYER INTERACTIONS

This paper presents briefly the recent progress on study of swept shock wave/boundary layer interactions with emphasis on application of zonal analysis and correlation analysis to them. Based on the zonal analysis an overall framework of complicated interaction flow structure including both surface flowfield and space flowfield is discussed. Based on correlation analysis the conical interactions induced by four families of shock wave generators have been discussed in detail. Some control parameter and physical mechanism of conical interaction have been revealed. Finally some aspects of the problem and the prospects for future work are suggested.

THE INTERACTION OF A SHOCK WAVE WITH THE BOUNDARY LAYER IN A REFLECTED SHOCK TUNNEL

The influence of a nontotal reflection on the interaction of a reflected shock wave with the boundary layer in a reflected shock tunnel has been investigated. The calculating method of the velocity, the temperature and the Mach number profiles in the boundary layer in reflected shock fixed coordinates has been obtained. To account for equilibrium real gas effects of nitrogen, the numerical results show that the minimum Mach number in the boundary layer has been moved from the wall into the boundary layer with the increasing of the incident shock Mach number. The minimum Mach number, the shock angle in the bifurcated foot and the jet velocity along the wall to the end plate are reduced owing to the Increasing of the area of nozzle throat. The numerical results are in good agreement with measurements.

Direct numerical simulation of shock wave/boundary layer interaction controlled by steady microjet in a compression ramp

Shock wave/boundary layer interaction in a 24°turning angle of the compression ramp at Mach number 2.9 controlled by steady microjet is investigated using direct numerical simulation.Three different jet spacings which are termed as sparse,moderate and dense are considered,and the induced vortex system and shock structures are compared.A moderate jet spacing configuration is found to generate counter-rotating vortex pairs that transport high-momentum fluid towards the vicinity of wall and strengthen the boundary layer to resist separation,reducing the separation region.The dense jet spacing configuration creates a larger momentum deficit region,reducing the friction downstream of the corner.Analysis of pressure and pressure gradient reveals that dense jet spacing configuration reduces the intensity of separation shock.The impact of varying jet spacings on the turbulent kinetic energy transport mechanism is also investigated by decomposing the budget terms in the transport equation.Furthermore,the spectral characteristics of the separation region are studied using power spectral density and dynamic mode decomposition methods,revealing that moderate jet spacing configuration suppresses low-frequency fluctuations in the separation region.

Effects of Number of Bleed Holes on Shock-Wave/Boundary-Layer Interactions in a Transonic Compressor Stator

An extensive numerical investigation is conducted to characterize the flow separation control in a transonic compressor cascade with a porous bleed.The bleed holes are arranged on the suction surface in a single row,two staggered rows and three staggered rows.For each bleed scheme,five bleed pressure ratios are examined at an inlet Mach number of 1.0.The results indicate that the aerodynamic performance of the cascade is significantly improved by the porous bleed.For the single-row scheme,the maximum reduction in total pressure losses is 57%.For the two-staggered-row and three-staggered-row schemes,there is an optimal bleed pressure ratio of 1.0,and the maximum reductions in total pressure loss are 68% and 75%,respectively.The low loss in the cascade is due to the well-controlled boundary layer.The new local supersonic region created by the bleed hole is the key reason for the improved boundary layer.The vortex induced by side bleeding provides another mechanism for delaying flow separation.Increasing the bleed holes could create multiple local supersonic regions,which reduce the range of the adverse pressure gradient that the boundary layer needs to withstand.This is the reason why cascades with more bleed holes perform better.

A CALCULATING METHOD OF SHOCK WAVE OSCILLATING FREQUENCY DUE TO TURBULENT SHEAR LAYER FLUCTUATIONS IN SUPERSONIC FLOW

One of the more severe fluctuating pressure environments encountered in supersonic or hypersonic flows is the shock wave oscillation driven by interaction of a shock wave with boundary layer. The high intensity oscillating shock wave may induce structure resonance of a high speed vehicle. The research for the shock oscillation used to adopt empirical or semiempirical methods because the phenomenon is very complex. In this paper a theoretical solution on shock oscillating frequency due to turbulent shear layer fluctuations has been obtained from basic conservation equations. Moreover, we have attained the regularity of the frequency of oscillating shock varying with incoming flow Much numbers M and turning angle . The calculating results indicate excellent agreement with measurements. This paper has supplied a valuable analytical method to study aeroelastic problems produced by shock wave oscillation.

Study on Internal Supersonic Flows with Pseudo-shock Wave Using Liquid Crystal Flow Visualization Method

The flow visualization technique using shear-sensitive liquid crystal is applied to the investigation of a Mach 2 internal supersonic flow with pseudo-shock wave （PSW） in a pressure-vacuum supersonic wind tunnel. It provides qualitative information mainly concerning the overall flow structure, such as the turbulent boundary layer separation, reattachment locations and the dimensionalities of the flow. Besides, it can also give understanding of the surface streamlines, vortices in separation region and the corner effect of duct flow. Two kinds of crystals with different viscosities are used in experiments to analyze the viscosity effect. Results are compared with schlieren picture, confirming the effectiveness of liquid crystal in flow-visualization.

Analysis of the Oscillatory Flows of Multiple Shock Waves in a Constant Area Duct

The oscillatory response of multiple shock waves to downstream perturbations in a supersonic flow is studied numerically in a rectangular duct.Multiple shock waves are formed inside the duct at a shock Mach number of 1.75.The duct has an exit height of H,and the effect of duct resonance on multiple shock oscillations is investigated by attaching exit ducts of lengths 0H,50H,and 150H.The downstream disturbance frequency varied from 10 Hz to 200 Hz to explore the oscillation characteristics of the multiple shock waves.The oscillatory response of shock waves under self-excited and forced oscillation conditions are analyzed in terms of wall static pressure,shock train leading-edge location,shock train length,and the size of the separation bubble.The extent of the initial shock location increases with an increase in exit duct length for the self-excited oscillation condition.The analysis of the shock train leading edge and the spectral analysis of wall static pressure variations are conducted.The variation in the shock train length is analyzed using the pressure ratio method for self-excited as well as forced oscillations.The RMS amplitude of the normalized shock train length(ζ_(ST))increases with an increase in the exit duct length for the self-excited oscillation condition.When the downstream perturbation frequency is increased,ζ_(ST)is decreased for exit duct configurations.For all exit duct designs and downstream forcing frequencies,the size of the separation bubble grows and shrinks during the shock oscillations,demonstrating the dependence on duct resonance and forced oscillations.

Hypersonic flow control of shock wave/turbulent boundary layer interactions using magnetohydrodynamic plasma actuators

The effect of magnetohydrodynamic(MHD)plasma actuators on the control of hypersonic shock wave/turbulent boundary layer interactions is investigated here using Reynolds-averaged Navier-Stokes calculations with low magnetic Reynolds number approximation.A Mach 5 oblique shock/turbulent boundary layer interaction was adopted as the basic configuration in this numerical study in order to assess the effects of flow control using different combinations of magnetic field and plasma.Results show that just the thermal effect of plasma under experimental actuator parameters has no significant impact on the flow field and can therefore be neglected.On the basis of the relative position of control area and separation point,MHD control can be divided into four types and so effects and mechanisms might be different.Amongst these,D-type control leads to the largest reduction in separation length using magnetically-accelerated plasma inside an isobaric dead-air region.A novel parameter for predicting the shock wave/turbulent boundary layer interaction control based on Lorentz force acceleration is then proposed and the controllability of MHD plasma actuators under different MHD interaction parameters is studied.The results of this study will be insightful for the further design of MHD control in hypersonic vehicle inlets.

Scaling of interaction lengths for hypersonic shock wave/turbulent boundary layer interactions

The interaction length induced by Shock Wave/Turbulent Boundary-Layer Interactions(SWTBLIs)in the hypersonic flow was investigated using a scaling analysis,in which the interaction length normalized by the displacement thickness of boundary layer was correlated with a corrected non-dimensional separation criterion across the interaction after accounting for the wall temperature effects.A large number of hypersonic SWTBLIs were compiled to examine the scaling analysis over a wide range of Mach numbers,Reynolds numbers,and wall temperatures.The results indicate that the hypersonic SWTBLIs with low Reynolds numbers collapse on the supersonic SWTBLIs,while the hypersonic cases with high Reynolds numbers show a more rapid growth of the interaction length than that with low Reynolds numbers.Thus,two scaling relationships are identified according to different Reynolds numbers for the hypersonic SWTBLIs.The scaling analysis provides valuable guidelines for engineering prediction of the interaction length,and thus,enriches the knowledge of hypersonic SWTBLIs.

Passage shock wave/boundary layer interaction control for transonic compressors using bumps

Flow separation due to shock wave/boundary layer interaction is dominated in blade passage with supersonic relative incoming flow,which always accompanies aerodynamic performance penalties.A loss reduction method for smearing the passage shock foot via Shock Control Bump(SCB)located on transonic compressor rotor blade suction side is implemented to shrink the region of boundary layer separation.The curved windward section of SCB with constant adverse pressure gradient is constructed ahead of passage shock-impingement point at design rotor speed of Rotor 37 to get the improved model.Numerical investigations on both two models have been conducted employing Reynolds-Averaged Navier-Stokes(RANS)method to reveal flow physics of SCB.Comparisons and analyses on simulation results have also been carried out,showing that passage shock foot of baseline is replaced with a family of compression waves and a weaker shock foot for moderate adverse pressure gradient as well as suppression of boundary layer separations and secondary flow of low-momentum fluid within boundary layer.It is found that adiabatic efficiency and total pressure ratio of improved blade exceeds those of baseline at 95%-100%design rotor speed,and then slightly worsens with decrease of rotatory speed till both equal below 60%rated speed.The investigated conclusion implies a potential promise for future practical applications of SCB in both transonic and supersonic compressors.

Unsteadiness of Shock Wave/Boundary Layer Interaction in Supersonic Cascade

An investigation of the passive control of shock wave/boundary laper interaction for reducing the amplitude of the shock oscillation was conducted on the circular arc-wedge (CW) profile cascade in a 220×290 mm transonic compressor cascade wind tunnel. A perforated surface with a cavity beneath it was positioned on the suction surface of the blade at the location of shock impingement. The Schlieren and high-speed photographs for flow over perforated blade are presented and compared with the results for solid blades. With the perforated surface, the high-speed photographs indicated an significant suppression of shock oscillation.

Shock Wave Smearing by Passive Control

Normal shock wave, terminating a local supersonic area on an airfoil, limits its performance and becomes a source of high speed impulsive noise. It is proposed to use passive control to disintegrate the shock wave. Details of the flow structure obtained by this method are studied numerically. A new boundary condition has been developed and the results of its application are verified against experiments in a nozzle flow. The method of shock wave disintegration has been confirmed and detailed analysis of the flow details is presented. The substitution of a shock wave by a gradual compression changes completely the source of the high speed impulsive noise and bears potential of its reduction.

Analysis of flow-field in a dual mode ramjet combustor with boundary layer bleed in isolator

A two-dimensional Reynolds averaged Navier Stokes(RANS)simulation of a dual mode ramjet(DMRJ)combustor is performed,modeling the University of Michigan dual-mode combustor experimental setup operating in reacting mode with different equivalence ratios(4).The simulations are carried out using a k-u SST turbulence model and a steady diffusion flamelet model for non-premixed combustion.Air enters the isolator at Mach 2.2,stagnation pressure and temperature of 549.2 kPa and 1400 K respectively.Hydrogen is injected transverse to the flow direction and upstream of the cavity flame holder to simulate ramjet(4 Z 0.29)and scramjet(4 Z 0.19)modes of operation.Wall static pressure plots are used to validate numerical results against experimental data.Analysis of flow separation in ramjet mode due to the presence of a shock train in the isolator is carried out by means of numerical Schlieren images overlapped with contours of negative axial velocity,showing the effects of shock wave boundary layer interaction(SWBLI).Active control through wall normal boundary layer bleed in the separated flow region is implemented,which weakens the shock train and moves it downstream closer to the cavity.Bleed results in an improved stagnation pressure recovery in ramjet mode,with a marginal increase in combustion efficiency.

Flow Control Effect on Unsteadiness of Shock Wave Induced Separation

In usual cases of significant pressure gradients and strong shocks, the front shock takes a fixed location along the wall, at which separation starts. Usually the rear shock is responding to vortex sheding by its deflection angle. In consequence main shock and rear shocks are moving whilst front shock is stable. The goal of the measurements presented here is to find out how the k-foot behaves during shock oscillations in the case when front shock is not fixed by the pressure gradient. Unsteady shock behaviour is also investigated when air jet vortex generators （AJVG） are used. Counteraction of the separation is directly related to the influence on unsteady processes in the shock wave induced separation.

Shock Wave Strength Reduction by Passive Control Using Perforated Plates

Strong, normal shock wave, terminating a local supersonic area on an airfoil, not only limits aerodynamic performance but also becomes a source of a high-speed impulsive helicopter noise. The application of a passive control system （a cavity covered by a perforated plate） on a rotor blade should reduce the noise created by a moving shock. This article covers the numerical implementation of the Bohning/Doerffer transpiration law into the SPARC code and includes an extended validation against the experimental data for relatively simple geometries of transonic nozzles. It is a first step towards a full simulation of a helicopter rotor equipped with a noise reducing passive control device in hover and in forward flight conditions.

NATURE OF THE SURFACE HEAT TRANSFER FLUCTUATION IN A HYPERSONIC SEPARATED TURBULENT FLOW

This paper presents the results of an experimental study of the unsteady nature of a hypersonic sepa- rated turbulent flow.The nominal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5x10^7/m.The separated flow was generated using finite span forward facing steps.An array of flush mounted high spatial resolution and fast response platinum film resistance thermometers was used to make mul- ti-channel measurements of the fluctuating surface heat trtansfer within the separated flow.Conditional sampling ana- lysis of the signals shows that the root of separation shock wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer thickness.The compression waves con- verge into a single leading shock beyond the boundary layer.The shock structure is unsteady and undergoes large-scale motion in the streamwise direction.The length scale of the motion is about 22 percent of the upstream influence length of the separation shock wave.There exists a wide band of frequency of oscillations of the shock system.Most of the frequencies are in the range of 1-3 kHz.The heat transfer fluctuates intermittently between the undisturbed level and the disturbed level within the range of motion of the separation shock wave.This inter mittent phenomenon is considered as the consequence of the large-scale shock system oscillations.Downstream of the range of shock wave motion there is a separated region where the flow experiences continuous compression and no intermittency phenomenon is observed.

Effects of Vortex Generator on Cylindrical Protrusion Aerodynamics

Experimental and numerical studies were carried out to evaluate the effect of vortex generator on a small cylindrical protrusion at Mach number 2.0. The experiments were performed using the supersonic blow down wind tunnel on different heights of cylindrical protrusion with vortex generator placed ahead of them. The upstream and downstream flow around the cylindrical protrusion is influenced by vortex generator as is observed using both visualization and pressure measurement techniques. Numerical studies using three dimensional steady implicit formulations with standard k-ω turbulence model was performed. Results obtained through the present computation are compared with the experimental results at Mach 2.0. Good agreements between computation and experimental results have been achieved. The results indicate that the aerodynamic drag acting on cylindrical protrusion can be reduced by adopting vortex generator.