Shock-induced Boundary Layer Separation Research Articles

Using the Reynolds Stress Model to Predict Shock-Induced Separation on Transport Aircraft

Additional results about the wing-body of the NASA Common Research Model at flight conditions derived from those of the AIAA Computational Fluid Dynamics Drag Prediction Workshop are introduced in this Paper. An initial assessment compared the widely used shear stress transport model with the full Reynolds stress model as two possible closures for the Reynolds-averaged Navier–Stokes problem. The study, based on the 6th AIAA Computational Fluid Dynamics Drag Prediction Workshop, confirmed the limitations of linear eddy-viscosity turbulence models in highly anisotropic turbulent flows. A side-of-body separation larger than previous experiments was predicted by the shear stress transport model. By comparison, the study revealed that the Reynolds stress model is capable of simulating higher-incidence transonic flight as the trends observed in the results were in overall good agreement with the experiment. The aim herein was to investigate the high-incidence transonic aerodynamics about the Common Research Model. Using Reynolds-Averaged Navier–Stokes closed with the Reynolds stress model, the aeroelastically deformed Common Research Model was then simulated at incidences outside of the Drag Prediction Workshop remit. The evolution of shock-induced boundary-layer separation was observed. A separation bubble was first predicted at an incidence of 3.5 deg, then developed in the chordwise and spanwise directions, triggering the stall of the aircraft at an incidence of 5.75 deg. Using shear lines helps one understand the three-dimensionality of this phenomenon as large areas of crossflow are present, starting in the wing-kink region and extending toward the wing tip.

Numerical investigation of nanosecond pulsed plasma actuators for control of shock-wave/boundary-layer separation

This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators that are designed to control shock-wave induced boundary-layer separation in a Mach 2.8 supersonic flow. By using two dielectric barrier surface discharge actuator configurations, parallel and canted with respect to the flow velocity vector, a previous experiment suggested that the actuator worked in two ways to influence the interaction: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the boundary-layer separation stronger, while the canted electrode produced vorticity and suppressed the boundary-layer separation due to the momentum transfer from the core flow. Because the detailed physical processes are still unclear, in this paper a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the plasma actuation, in which the dielectric barrier discharge produced plasma is approximated as a high temperature region. The flow characteristics without the plasma actuation correspond to the experimental observation, indicating that the numerical method successfully resolves the shock-wave/boundary-layer interaction. With the plasma actuation, complete agreement between the experiment and calculation has not been obtained in the size of the shock-wave/boundary-layer interaction region. Nevertheless, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the strength of the interaction, while the canted electrode leads to a reduction of the interaction strength, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the shock induced boundary layer separation.

Confinement effects on regular–irregular transition in shock-wave–boundary-layer interactions

An oblique shock wave is generated in a Mach 2 flow at a flow deflection angle of$12^{\circ }$. The resulting shock-wave–boundary-layer interaction (SWBLI) at the tunnel wall is observed. A novel traversable shock generator allows the position of the SWBLI to be varied relative to a downstream expansion fan. The relationship between the SWBLI, the expansion fan and the wind tunnel arrangement is studied. Schlieren photography, surface oil flow visualisation, particle image velocimetry and high-spatial-resolution wall pressure measurements are used to investigate the flow. It is observed that stream-normal movement of the shock generator downwards (towards the floor and hence the point of shock reflection) is accompanied by (1) growth in the streamwise extent of the shock-induced boundary layer separation, (2) upstream movement of the shock-induced separation point while the reattachment point remains nearly fixed, (3) an increase in separation shock strength and (4) transition between regular and irregular (Mach) reflection without an increase in incident shock strength. The role of free interaction theory in defining the separation shock angle is considered and shown to be consistent with the present measurements over a short streamwise extent. An SWBLI representation is proposed and reasoned which explains the apparent increase in separation shock strength that occurs without an increase in incident shock strength.

Control of Incident Shock-Induced Separation Using Vane-Type Vortex-Generating Devices

An experimental investigation was conducted to control an incident shock-induced boundary-layer separation associated with a 14 deg shock generator in a Mach 2.05 flow. Two vane-type configurations, namely the triangular (, 0.5, 0.8, 1.0) and the rectangular () designs, were studied. An array of each control device was tested for three control locations of , 10, and 15. The control location of is seen to show the maximum reduction in separation length for each device tested. For the rectangular-vane device (), a maximum reduction of 38% in separation length is observed, followed by the triangular-vane devices of and 1.0, each of which shows a 32% reduction, and finally, with 18%. The effectiveness of these devices to control separation is, however, seen to decrease with increase in . In terms of separation shock unsteadiness, the maximum rms value for shows the highest value for each control device, and this value decreases with increase in control location. At , both the rectangular vane () and triangular vane () show a 50% reduction in maximum rms value, whereas it decreases to 30% at for these devices.

Control of incident shock-induced boundary-layer separation using steady micro-jet actuators at M∞ = 3.5

Experiments were carried out to control an incident shock-induced separation associated with 22° shock generator in a Mach 3.5 flow using an array of steady micro-jet actuators. Four micro-jet actuator configurations based on the variation in their pitch angle [Formula: see text], skew angle [Formula: see text] and span-wise spacing were used. Each of these configurations were placed 14 δ upstream of the interaction and operated with injection pressures ( Poj) varying from 140 to 643 kPa. While no major variations in separation characteristics were observed for Poj < 140 kPa, significant modifications were observed beyond [Formula: see text] of 140 kPa and until 208.5 kPa. Amongst all the four control configurations, micro-jet vortex generator 2 ([Formula: see text] showed the best control with a 2 δ downstream shift in separation point location relative to no-control. The shift is also accompanied with a change in maximum zero-crossing frequency towards higher frequency (almost twice), a reduction in the intermittency length and an increase in the correlation value between the boundary layer just upstream of the interaction and the intermittent region. These results indicate that the effectiveness of micro-jet vortex generator 2 is probably due to the improved entrainment levels in the shear layer induced by the micro-vortices which are generated downstream of these devices. The increase of the skew angle [Formula: see text] from 180° to 270° for the same pitch angle of β = 45° (micro-jet vortex generator 3) seems to have no major impact on the separation characteristics. The reduction in the span-wise spacing (micro-jet vortex generator 4) resulted in deterioration of the flow field due to the jet-to-jet interaction with increasing injection pressures.

Structural characteristics of the shock-induced boundary layer separation extended to the leading edge

For a better understanding of the local unstart of supersonic/hypersonic inlet, a series of experiments has been conducted to investigate the shock-induced boundary layer separation extended to the leading edge. Using the nanoparticle-based planar laser scattering, we recorded the fine structures of these interactions under different conditions and paid more attention to their structural characteristics. According to their features, these interactions could be divided into four types. Specifically, Type A wave pattern is similar to the classic shock wave/turbulent boundary layer interaction, and Type B wave configuration consists of an overall Mach reflection above the large scale separation bubble. Due to the gradual decrease in the size of the separation bubble, the separation bubble was replaced by several vortices (Type C wave pattern). Besides, for Type D wave configuration which exists in the local unstart inlet, there appears to be some flow spillage around the leading edge.

Multidimensional High-Resolution Schemes for Viscous Hypersonic Flows

The first-order multidimensional flux difference splitting method (Parent, B., “Multidimensional Flux Difference Splitting Schemes,” AIAA Journal, Vol. 53, No. 7, 2015, pp. 1936–1948) is here extended to second-order accuracy through the use of limiters. The proposed second-order flux functions are such that they collapse to the first-order multidimensional flux difference splitting stencil in the vicinity of discontinuities and tend toward second-order finite difference stencils in smooth parts of the solution. Various test cases of interest to hypersonic flight including supersonic ramp injectors, shock-induced boundary-layer separation, and hypersonic viscous layers reveal that the proposed method achieves as much as a four times increase in resolution per dimension over its first-order counterpart or over the second-order symmetric total variation diminishing schemes. When integrated through a block-implicit approach, this entails as much as a 40-fold increase in computational efficiency while not compromising the symmetric total variation diminishing reliable convergence to steady-state, essentially monotone solution, and high resolution of boundary layers.

Flow visualization for shock-induced boundary-layer separation extended to the flat-plate-leading edge

Flow visualization for shock-induced boundary-layer separation extended to the flat-plate-leading edge

Unsteady Operation of a Highly Supersonic Organic Rankine Cycle Turbine

Organic Rankine cycle (ORC) turbogenerators are the most viable option to convert sustainable energy sources in the low-to-medium power output range (from tens of kWe to several MWe). The design of efficient ORC turbines is particularly challenging due to their inherent unsteady nature (high expansion ratios and low speed of sound of organic compounds) and to the fact that the expansion encompasses thermodynamic states in the dense vapor region, where the ideal gas assumption does not hold. This work investigates the unsteady nonideal fluid dynamics and performance of a high expansion ratio ORC turbine by means of detailed Reynolds-averaged Navier–Stokes (RANS) simulations. The complex shock interactions resulting from the supersonic flow (M ≈ 2.8 at the vanes exit) are related to the blade loading, which can fluctuate up to 60% of the time-averaged value. A detailed loss analysis shows that shock-induced boundary layer separation on the suction side of the rotor blades is responsible for most of the losses in the rotor, and that further significant contributions are given by the boundary layer in the diverging part of the stator and by trailing edge losses. Efficiency loss due to unsteady interactions is quantified in 1.4% in absolute percentage points at design rotational speed. Thermophysical properties are found to feature large variations due to temperature even after the strong expansion in the nozzle vanes, thus supporting the use of accurate fluid models in the whole turbine stage.

Computational study and error analysis of an integrated sampling-probe and gas-analyzer for mixing measurements in supersonic flow

Concentration probes are employed in supersonic flow mixing measurements. Because the typical design of such probes is essentially based on an inviscid, adiabatic, quasi-1D analysis, the scope of this work is to understand better and quantify the severe impact of viscous effects on the probe’s internal gasdynamics and the associated uncertainties in the measured quantities via a computational fluid dynamics analysis. Specifically, the focus is on the augmented errors due to the aforementioned viscous effects when coupled with various cases of probe-flow misalignment, which is a typical scenario encountered in mixing measurements of binary gas compositions (air and helium in the present work) in vortex-dominated flows. Results show phenomena such as shock induced boundary layer separation and the formation of an oblique shock train. These flow features are found to noticeably affect the accuracy of the composition measurement. The errors associated with the inviscid, adiabatic, quasi-1D analysis of the probes are quantified in this study.

Shock wave interactions between slender bodies

A complex interference flowfield consisting of multiple shocks and expansion waves is produced when high-speed slender bodies are placed in close proximity. The disturbances originating from a generator body impinge onto the adjacent receiver body, modifying the local flow conditions over the receiver. This paper aims to uncover the basic gas dynamics produced by two closely spaced slender bodies in a supersonic freestream. Experiments and numerical simulations were used to interpret the flowfield, where good agreement between the predictions and measurements was observed. The numerical data were then used to characterise the attenuation associated with shock wave diffraction, which was found to be interdependent with the bow shock contact perimeter over the receiver bodies. Shock-induced boundary layer separation was observed over the conical and hemispherical receiver bodies. These strong viscous-shock interactions result in double-reflected, as well as double-diffracted shock wave geometries in the interference region, and the diffracting waves progress over the conical and hemispherical receivers’ surfaces in “lambda” type configurations. This gives evidence that viscous effects can have a substantial influence on the local bow shock structure surrounding high-speed slender bodies in close proximity.

Investigation of shock-induced boundary layer separation extended to the flat plate leading-edge

Experimental investigations and numerical simulations have been carried out to investigate the shock-induced boundary layer separation, which is extended to the flat plate leading-edge. Three typical separation structures are obtained when an incident shock interacts with an boundary layer: one contains an overall regular reflection, one contains an overall Mach reflection, and one contains a detached shock. All the three structures show comparatively stable characteristic in the experiment. Numerical results demonstrate that there possibly exist two different flow field solutions for a uniform flow condition from different initializations. It could be explained as the asymmetric shock reflection hysteresis phenomenon, analyzed from the shock polars.

Shock-Induced Boundary-Layer Separation in Round Convergent–Divergent Nozzles

The location of shock-induced boundary-layer separation inside the divergent section of convergent–divergent nozzles is studied experimentally. Pressure-sensitive paint technique is used with nozzles of different design Mach numbers in the range 1.4–2.8. Nozzle pressure ratios in the range of 1.12–4.91, corresponding to a “jet Mach number” range of 0.4–1.7 and a Reynolds number range of –, are covered in the experiment. As it is well-known, one-dimensional nozzle flow theory grossly overpredicts the throat-to-shock-location distance at a given nozzle pressure ratio. A correlation from the literature based on rocket nozzle databases is also found to be inadequate for these nozzles of lower design Mach number typical of aircraft applications. For the parametric range covered, a simple correlation for the shock location distance is found. All data collapse in a cluster when plotted as a function of the ratio of jet Mach number to design Mach number. A curve-fit equation representing the average trend is provided.

A review of the use of vortex generators for mitigating shock-induced separation

This article reviews research into the potential of vortex generators to mitigate shock-induced separation. Studies ranging from those conducted in the early post-war era to those performed recently are discussed. On the basis of the investigations described in this report, it is clear that vortex generators can alleviate shock-induced boundary layer separation. Yet, it will be shown that their potential and efficiency varies considerably in practical applications. Much more success is reported in transonic test cases compared to separation induced in purely supersonic interactions. Under a variety of flow conditions, the best performance is achieved with vortex generators with a height of roughly half the boundary layer thickness and a shape similar to a swept vane. Notwithstanding this, vortex generator performance is not as consistent as it is in low-speed applications. Further work is required before vortex generators can be implemented into the design process for eliminating shock-induced separation on transonic wings and in supersonic inlets.

Numerical Investigation into the Combustion Behavior of an Inlet-Fueled Thermal-Compression-Like Scramjet

A numerical study on the combustion behavior of an inlet-fueled three-dimensional nonuniform-compression scramjet is presented. This paper is an extension to previous work on the combustion processes in a premixed three-dimensional nonuniform-compression scramjet, where thermal compression was shown to enhance combustion. This paper demonstrates how thermal compression can be used in a generic scramjet configuration with a realistic fuel-injection method to enhance performance at high flight Mach numbers. Such a scramjet offers an extra degree of freedom in the design process of fixed-geometry scramjets that must operate over a range of flight Mach numbers. In this study, how the combustion processes are affected is investigated, with the added realism of inlet porthole fuel injection. Ignition is established from within a shock-induced boundary-layer separation at the entrance to the combustor. Radicals that form upstream of the combustor within the inlet, from the injection method, enhance combustion. Coupling of the inlet-induced spanwise gradients and thermal compression improves combustion. The results highlight that, although the fuel-injection method imparts local changes to the flow structures, the global flow behavior does not change compared to previous premixed results. This combustion behavior will be reproduced when using other fueling methods that deliver partially premixed fuel and air to the combustor entrance.

Passive Control of Shock Oscillation Around a Biconvex Circular Arc Airfoil in a Channel

A strong normal shock wave, generated on an airfoil, is responsible not only for limiting the aerodynamic performance but also for shock induced boundary layer separation. This shock induced boundary layer separation results aerodynamics instabilities (buffet), high cycle fatigue failure (HCF), nonsynchronous vibration (NSV), flutter and so on. In the present study, a numerical computation has been performed to control the unsteady shock oscillation over a 12% biconvex circular arc airfoil in a two dimensional channel. Reynolds averaged Navier-Stokes (RANS) equations with k-ω shear stress transport (SST) two equation turbulence model has been applied for computational analysis. To control the shock oscillation over a biconvex circular arc airfoil (referred as base airfoil), the geometry of the base airfoil has been modified by incorporating a cavity with two openings on both upper and lower surface of the airfoil. The cavity has been incorporated in such a manner that the mean position (along chord length) of the cavity is placed where the RMS of static pressure fluctuation on airfoil surface (for base airfoil) is maximum. The length and depth of the cavities are kept 10% and 2% of the chord length, respectively. The behavior of the shock wave oscillation has been studied for a particular pressure ratio (defined as the ratio of back pressure to inlet total pressure) of 0.69. The present study investigates the shock wave characteristics over (a) airfoil with no cavity (base airfoil) and (b) airfoil with cavity with 60% opening (60% of cavity length is open). The results showed that the incorporation of cavities on airfoil surfaces not only affect the flow field but also change the behaviour in a great extent. For pressure ratio 0.69, the flow field becomes steady for airfoil with cavities while for base airfoil the flow field was unsteady. The results also show that incorporation of a cavity on airfoil surfaces changes the type of shock wave from normal to λ shock wave.

Numerical Simulation on the Shock Wave/Boundary Layer Interaction with Heating/Cooling Effect

The complicated flow conditions inner or outer of the high speed vehicle would make the shock wave boundary layer (SWBLI) much different from its standard state on an adiabatic wall. The heating and cooling effects on the SWBLI achieved by changing the wall temperature, are evaluated by the numerical simulations. It is found that the wall temperatures have significant influence on the flow patterns of SWBLI. The scale of shock induced boundary layer separation increases obviously with the rise of the wall temperature, and the separation would shrink dramatically with the decrease of the wall temperature. This could be partially explained by the variation of the momentum thickness and displacement thickness of the incoming turbulent boundary layer under different wall temperature.

Effect of Reynolds number on compressible convex-corner flows

An experimental study was conducted to investigate the effect of Reynolds number on compressible convex-corner flows, which correspond to an upper surface of a deflected flap of an aircraft wing. The flow is naturally developed along a flat plate with two different lengths, resulting in different incoming boundary layer thicknesses or Reynolds numbers. It is found that boundary layer Reynolds number, ranging from <TEX>$8.04{\times}10^4$</TEX> to <TEX>$1.63{\times}10^5$</TEX>, has a minor influence on flow expansion and compression near the corner apex in the transonic flow regime, but not for the subsonic expansion flow. For shock-induced separated flow, higher peak pressure fluctuations are observed at smaller Reynolds number, corresponding to the excursion phenomena and the shorter region of shock-induced boundary layer separation. An explicit correlation of separation length with deflection angle is also presented.

Supersonic Combustion Processes in a Premixed Three-Dimensional Nonuniform-Compression Scramjet Engine

A numerical study was undertaken using the commercial computational-fluid-dynamics code to analyze the combustion behavior of a three-dimensional nonuniform-compression scramjet. Work here was inspired by Ferri’s concept of thermal compression and extended this to include the influence of three-dimensional viscous flow phenomena on the combustion behavior. A premixed mixture with an equivalence ratio of 1 was used to eliminate the influence of the fuel-injection method on the combustion characteristics. The freestream properties corresponded to a Mach 10 flight condition that gave a flow enthalpy, total temperature, and total pressure of , 4676 K, and 29 MPa, respectively. The combustion flame was found to propagate throughout the entire bottom wall of the combustor. Three flame propagation processes were identified: three-dimensional flow structures that provide ignition sources within the boundary layer, radical transport within a three-dimensional shock-induced boundary-layer separation, and thermal compression due to combustion within the high compression region. The investigation successfully characterises each individual flame propagation mechanism on the flowfield. The results highlight the closely coupled interaction between the fluid dynamic and combustion processes within a three-dimensional scramjet engine, which offers insight into how these mechanisms can be exploited to provide superior scramjet performance over a range of flight Mach numbers.

An experimental study of transonic shock/boundary layer interactions subject to downstream pressure perturbations

An experimental study of the response to downstream pressure perturbations of transonic shocks in a parallel walled duct has been conducted. Tests have been carried out for normal shock strengths of M ∞ = 1.4 and 1.5 with pressure perturbation frequencies in the range 16–90 Hz. The measurement techniques of high speed schlieren photography, laser Doppler anemometry and high frequency pressure measurements have been used. In all test cases, the normal shock undergoes periodic oscillatory motion in the streamwise direction. The interaction structure between the oscillating shock and the tunnel wall turbulent boundary layer varies during oscillations, especially at M ∞ = 1.4 , where the leading shock–leg of the lambda shock–foot structure gets significantly stronger during upstream shock motion and weaker during downstream motion. These changes are related to the relative strength of the shock (which varies due to the velocity of shock motion) and the effect that this has on the extent of boundary layer thickening and/or shock induced boundary layer separation. The dominant mechanism that governs the dynamics of shock motion in response to downstream pressure perturbations has also been identified and explains why the amplitude of shock motion decreases with increasing frequency.