Shock-induced Separation Research Articles

High-speed microactuation in a supersonic dual-stream jet flow

Supersonic shear layers experience instabilities that generate significant adverse effects; in complex configurations, these instabilities have global impacts as they foster compounding complications with other independent flow features. We consider the flow near the exit of a dual-stream rectangular nozzle. The supersonic core and sonic bypass streams mix downstream of a splitter plate trailing edge (SPTE) just above an adjacent deck. We perform two-dimensional direct numerical simulations with laminar inflow conditions to parametrically explore the influence of active flow control, considering various actuation angles and locations. The goal is to alleviate the prominent tone associated with vortices shed at the SPTE; these vortices initiate an unsteady shock system that affects the entire flow field through a shock-induced separation and the downstream evolution of plume shear layers. Resolvent analysis is performed on the baseline flow. The identified optimal response guides the placement of steady-blowing actuators. Since the resolvent analysis fundamentally investigates the input–output dynamics of a system, it is also utilized to uncover actuation-induced changes in the forcing–response dynamics. Spectral analysis shows that the baseline flow fluctuating energy is concentrated in the shedding instability. Actuating at optimal angles based on location disperses this energy into various flow features; this affects the shedding itself, and the structure and unsteadiness of the shock system, and thus the response of the deck and nozzle wall boundary layers and the plume. The resolvent analysis indicates, and Navier–Stokes solutions confirm, that favourable control is obtained by either indirectly or directly mitigating the baseline instability.

Highly separated transitional shock-wave/boundary-layer interactions: A spatial modal study

At cruise altitudes, the Reynolds number may become sufficiently low to allow a laminar boundary layer to persist on the suction side of a transonic fan blade up to the shock-wave/boundary-layer interaction (SBLI). In such a transitional SBLI with sufficiently large shock-induced separation, a shock oscillation mechanism occurs, with the source at the upstream growth (until a critical length) and natural suppression (through shear layer instabilities) of the separation bubble. The oscillation cycle is characterized by a temporarily vanishing upstream laminar part of the separation bubble. The suppression of this laminar part is accompanied by downstream advection of turbulence and subsequent entrainment into the bulk separation bubble, affecting the reflected shock movement. In order to study the spatial behavior of the mechanism, particle image velocimetry of a highly separated transitional oblique SBLI at Mach 2.3 is conducted in the high speed aerodynamics laboratory of Delft University of Technology. Statistical quantities, including root mean square velocity fluctuations, phase averages, and spatial modes from proper orthogonal decomposition, are investigated. The entrainment strength varies depending on the phase of the oscillation, and the turbulent shear layer is not fully developed. The main growth and shrinking mode of the separation bubble were extracted, which affects the slip line region size and shock position. Secondary modes that affect the shear layer undulation and upstream effects were also extracted. The study provides quantitative analyses of an important shock oscillation type, with the focus on capturing the separation bubble size variation and upstream effects.

Point-enhanced convolutional neural network: A novel deep learning method for transonic wall-bounded flows

Low order models can be used to accelerate engineering design processes. Ideally, these surrogates should meet the conflicting requirements of large design space coverage, high accuracy and fast evaluation. Within the context of aerospace applications at transonic conditions, this can be challenging due to the associated non-linearity of the flow regime. Different methods have been investigated in the past to predict the flow-field around shapes such as airfoils or cylinders. However, they usually have reduced spatial resolution, limiting the prediction capabilities within the boundary layer which is of interest for transonic wall-bounded flows. This work proposes a novel Point-Enhanced Convolutional Neural Network (PCNN) method that combines the advantages of the well-established PointNet and convolutional neural network approaches. The PCNN model has relatively low memory requirements in the training process, preserves the spatial correlation in the domain and has the same resolution as a traditional computational method. The architecture is used for the flow-field prediction of civil aero-engine nacelles in which it is demonstrated that the flow features of peak isentropic Mach number (Mis), pre-shock isentropic Mach number and shock location (X/Lnac) are captured within ΔMis = 0.02, ΔMis=0.04, ΔX/Lnac=0.007, respectively. The PCNN model successfully predicts the integral parameters of the boundary layer, in which the incompressible displacement thickness, momentum thickness and shape factor are typically within 5% of the CFD. Overall, the PCNN method is demonstrated for transonic wall-bounded flows for a range of flow physics that include shock waves and shock-induced separation.

Optimization design method based on parameter reduction and active subspaces: Redistribution of chordwise loading at blade tips in a transonic axial-flow fan

In practical optimization design, an excessive number of design variables have a highly detrimental influence on the efficiency and accuracy of the final design scheme and expose the optimization problem to the curse of dimensionality. Therefore, incorporating only the most essential variables into an optimization design problem facilitates obtaining accurate and cost-efficient solutions. Reported here is an optimization design method based on parameter reduction and active subspaces, and it is used to redistribute the tip load in a transonic fan. Specifically, a coupled design strategy is developed to reduce the number of parameters needed to describe the three-dimensional blade shape, which leads to far fewer design variables being involved in the optimization design. Moreover, active subspaces are used to perform sensitivity analysis and establish low-dimensional surrogate models. After the coupled design, a blade is represented effectively by only three parameters, each of which has a significant influence on the fan performance. Three one-dimensional active subspaces are established for maximum mass flow rate, maximum total pressure ratio, and maximum efficiency, based on which the linear surrogate models are obtained. Next, the chordwise tip blade loading is optimized, after which the rotor efficiency at the design point is increased by 1.1%, while the total pressure ratio remains nearly unchanged. Finally, the flow field is analyzed to understand the mechanism for this performance improvement, and the results show that the optimized blade loading reduces the aerodynamic losses caused by shock-induced flow separation and the interaction between shocks and tip leakage flows.

Pressure feedback system for flow separation mitigation in scramjet intakes

Shock Wave Boundary Layer Interactions (SWBLI) occur due to the convergence of a shock wave and a viscous boundary layer. The resulting separation bubble is a major setback for the performance of high-speed air intakes. This paper presents a numerical investigation into the potential of a Pressure Feedback Technique (PFT) to alleviate shock-induced flow separation within a Scramjet engine intake operating at Mach 4. The PFT is a novel self-sustaining flow control technique that combines simultaneous suction and injection. The two main output parameters used to support the hypothesis of the present study are the size of the separation bubble as well as the total pressure recovery at the isolator outlet. The most prominent observation of this study is that with the installation of the PF tubes, the total pressure recovery at the exit is noted to rise. The effectiveness of the pressure feedback technique in reducing the intensity of the separation bubbles is found to depend on the diameter of the PFT, pitch-to-diameter ratio, and PFT tube design.

Numerical investigations of shock/boundary-layer interaction and control for Mach 2.5 flow in an axisymmetric inlet

The present work investigates the application of vortex generators for separation control in axisymmetric isolator flows. Reynolds-Averaged Navier-Stokes computations were performed to simulate a Mach 2.5 flow past a half-isolator geometry with a 20∘ axisymmteric compression ramp based on the experiments of Funderburk and Narayanaswamy at North Carolina State University. Single vortex generator and a vortex generator array placed upstream of the shock-induced separation region were investigated. Near wall streamlines, surface pressure contours and density contours on crosswise planes were compared with experiments for flow control using a single vortex generator. Results indicate that the present computations are able to capture the wake and shock structures, and also predict reduction in the streamwise extent of flow separation downstream of the device trailing edge. Finally, the effects of the shape/orientation (forward facing/backward facing) of a single vortex generator, and the design of an array of three vortex generators (with devices of similar and mixed orientations) on the flow separation were also investigated. Contours of near-surface streamwise velocity showed that device orientation had a strong effect on separation control, which is attributed to the differences in the primary streamwise vortices shed in the two cases. Further, the overall reduction in the footprint of separated flow was determined to be most with the use of an array of vortex generators with mixed orientations.

Interaction Length Analysis of Hypersonic Shock–Boundary-Layer Interaction Including High-Enthalpy Effects

Shock-induced flow separation plays an important role in practical applications of hypersonic flows, like intake unstart and unsteady aeroelastic phenomena. The size of the shock–boundary-layer interaction (SBLI) region determines the extent of choking in engine inlet ducts. Recent scaling studies have shown that interaction length is a function of shock-induced pressure jump, upstream Mach number, Reynolds number, wall heating/cooling, and flow deflection angle at the shock wave. In this work, we analyze the interaction length scaling for different hypersonic SBLI cases at varying flow enthalpies. We perform Reynolds-averaged Navier–Stokes (RANS) simulations, using a shock-strength-dependent turbulence model, of hypersonic SBLI flows over a range of Mach numbers and at flow enthalpies where vibrational nonequilibrium starts to be important. The numerical results closely match the experimental data, validating the RANS predictions. The size of the separation bubble in such high-enthalpy cases is found to not follow the known scaling relations. We propose a modified scaling that collapses a wide range of hypersonic Mach numbers and freestream enthalpy conditions. We also derive the high-Mach-number limit to demonstrate that wall cooling/heating and geometry are the main factors affecting the size of the SBLI region in hypersonic flows.

Improved Effectiveness of Vane-Type Vortex Generators for Incident Shock-Induced Separation

An experimental investigation is conducted to improve the control effectiveness of an array of vane-type vortex generators (VG) implemented 6 δ upstream of an incident shock-induced separation generated using a 14-deg wedge in a Mach 2.05 flow. Initially, the effect of an array of rectangular vanes (RRV) is studied by varying the vane 1) chord to height (c/h)=7.2, 4.2, 2) angle α=24, 20, 18, and 16 deg, 3) height to boundary-layer thickness (h/δ)=0.75, 0.5, 0.3, and 0.2, and 4) interdevice spacing to height (s/h)=12, 9.5, 7.5, 6.5, 5.7, 5.5, and 5.0. Similar tests were also performed for ramp vanes (RV) with α=24 deg. Due to inherent design differences, an RRV device generates a larger region of vortex influence relative to an RV device. Therefore, reducing s/h significantly reduces the extent of uncontrolled separation regions between neighboring VGs that improves their separation control capability. The best performing RRV devices with vane height h/δ=0.5, h/δ=0.3, and h/δ=0.2 show maximum reduction in separation extent of nearly 80, 74, and 70%, respectively, relative to no control. The equivalent RV devices, however, show much reduced control effectiveness. All controls exhibit a shift in the dominant separation frequency to higher values, indicating a reduction in the overall extent of separation.

Simulation of Hypersonic Shock-Boundary Layer Interaction Using Shock-Strength Dependent Turbulence Model

Shock-induced flow separation is important in hypersonic intake unstart and for unsteady aero-thermal loads. This work presents a computational study of shock-boundary layer interaction (SBLI) at hypersonic Mach numbers. The objective is to accurately predict the separation bubble size for oblique shock impingement on a turbulent boundary layer and for two-dimensional axisymmetric cone-flare SBLI cases. We use the Reynolds-averaged Navier Stokes method with an advanced shock-strength-dependent turbulence model. A recently developed transport equation-based shock sensor is employed to identify the location and strength of shock waves. The computational results are found to match experimental measurements of surface pressure and skin-friction coefficient for a series of test cases, highlighting the effects of Mach number, shock generator angle, and wall cooling on the size of the separation bubble. Results are also found to conform to the scaling of SBLI separation size found in the literature. This is one of the few such corroborations for hypersonic axisymmetric SBLI.

Thermal optimization of shock-induced separation in a natural laminar airfoil operating at off-design conditions

Thermal optimization of shock-induced separation in a natural laminar airfoil operating at off-design conditions

Design considerations for efficient spanwise-inclined air-jet vortex generators for separation control in supersonic and hypersonic flows

Shock-wave/boundary-layer interactions and associated flow separation are frequently occurring phenomena in many high-speed flow applications. Injection of spanwise-inclined air-jets is an efficient means to control shock-induced flow separation in supersonic and transonic flows. Separation-control studies in hypersonic flow are scarce and so far mostly restricted to microramp control. The objectives of the current work are to broaden the scope into hypersonic regime, investigate the influence of crossflow Mach and Reynolds number on the jet-injection flow field, and to find a suitable control parameter that is valid across a wide range of flow conditions. For this purpose, we study injection of a single spanwise-inclined jet in supersonic and hypersonic crossflows using large-eddy simulations. We analyzed the flow topology, induced vortical structures and boundary-layer statistics. The general characteristics of the jet/crossflow interactions are similar in the supersonic and hyperonic regimes: the injection of a spanwise-inclined jet induces an asymmetric flow topology with a complex system of shocks and vortical structures. We identified the ratio of injection-pressure-to-freestream-pressure as a suitable parameter to characterize and compare jet/crossflow interactions in different flow regimes. In addition, we recommend larger boundary-layer-to-jet-diameter ratio for control applications, because it delays the lift-off of major vortices. The present results suggest that rows of spanwise-inclined jets can be efficiently used to control separation also in hypersonic flow. Moreover, low momentum-flux jet injection can be potentially also used for cooling purposes.

Mitigation of Shock-Induced Separation Using Square-Shaped Micro-Serrations—A Preliminary Study

Suppressing shock-induced flow separation has been a long-standing problem in the design of supersonic vehicles. To reduce the structural and design complexity of control devices, a passive control technique based on micro-serrations is proposed and its controlling effects are preliminarily investigated under test conditions in which the Mach number is 2.5 and the ramp creating an incident shock is 15 deg. Meanwhile, a vorticity-based criterion for assessing separation scales is developed to resolve the inapplicability of the zero skin friction criterion caused by wall unevenness. The simulations demonstrate that the height of the first stair significantly influences the separation length. Generally, the separation length is shorter at higher stairs, but when the height is greater than half of the thickness of the incoming boundary layer, the corresponding separation point moves upstream. A stair with a height of only 0.4 times the thickness of the boundary layer reduces the separation length by 2.69%. Further parametric analysis reveals that while the remaining serrations have limited effects on the flow separation, an optimization of their shape (depth and width) can create more favorable spanwise vortices and offer a modest improvement of the overall controlling performance. Compared to the plate case, a 9.13% reduction in the separation length can be achieved using a slightly serrated design in which the leading stair is 0.1 high and the subsequent serrations are 0.2 deep and 0.05 wide (nondimensionalized, with the thickness of the incoming boundary layer). Meanwhile, the micro-serration structure even brings less drag. Considering the minor modification to the structure, the proposed method has the potential for use in conjunction with other techniques to exert enhanced control on separations.

Application of Hybrid RANS/LES Methods for the Simulation of Shock-Induced Turbulent Boundary Layer Separation

Application of Hybrid RANS/LES Methods for the Simulation of Shock-Induced Turbulent Boundary Layer Separation

Shock-Induced Separation and Control in a 4.32-Aspect-Ratio Test Section

Experimental test sections studying shock-wave–boundary-layer interactions are usually right-angled by design to replicate the supersonic engine inlets where these interactions are observed. To avoid flow separation and shock effects on the side walls, investigations are typically performed on the centerline of the test section. However, the centerlines of small-aspect-ratio test sections can be significantly affected by the flow structures generated in the corners. Results using a large-aspect-ratio test section at a freestream Mach number of 1.6 include centerline and corner flow separation topologies, schlieren images of shock structures, total pressure profiles, and surface static pressures throughout the interaction region. These results provide evidence that the large-aspect-ratio test section can produce a large area of centerline flow that is unaffected by the separation and shock structures generated in the corners, which is a significant improvement over most of the smaller-aspect-ratio test sections employed for these types of studies. Microvortex generators are also employed upstream of the interaction, and measurements indicate a five-vortex system that energizes the near-wall boundary layer enough to overcome the shock’s adverse pressure gradient while remaining attached.

Flow Structure behind Spanwise Pin Array in Supersonic Flow

This work focused on the experimental characterization of a complex flow structure behind a cross-flow array of cylindrical pins installed on the wall of a supersonic duct. This geometry simulates several common gas dynamic configurations, such as a supersonic mixer, a turbulence-generating grid, or, to some extent, a grid fin. In this work, the instrumentation employed is essentially non-intrusive, including spanwise integrating techniques such as (1) fast schlieren visualization and (2) Shack–Hartmann wavefront sensors; and planar techniques, namely (3) acetone Mie scattering and (4) acetone planar laser-induced fluorescence. An analysis of the data acquired by these complementary methods allowed the reconstruction of a three-dimensional portrait of supersonic flow interactions with a discrete pin array, including the shock wave structure, forefront separation zone, shock-induced separation zone, shear layer, and the mixing zone behind the pins. The main objective of this activity was to use various visualization techniques to acquire essential details of a complex compressible flow in a wide range of temporal–spatial scales. Particularly, a fine structure in the supersonic shear layer generated by the pin tips was captured by a Mie scattering technique. Based on the available publications, such structures have not been previously identified or discussed. Another potential outcome of this work is that the details revealed could be utilized for adequate code validation in numerical simulations.

Moving wall effect on normal shock wave–turbulent boundary layer interaction on an airfoil

PurposeThis paper aims to describe a proposal for an innovative method of normal shock wave–turbulent boundary layer interaction (SBLI) and shock-induced separation control.Design/methodology/approachThe concept is based on the introduction of a tangentially moving wall upstream of the shock wave and in the interaction region. The SBLI control mechanism may be implemented as a closed belt floating on an air cushion, sliding over two cylinders and forming the outer skin of the suction side of the airfoil. The presented exploratory numerical study is conducted with SPARC solver (steady 2D RANS). The effect of the moving wall is presented for the NACA 0012 airfoil operating in transonic conditions.FindingsTo assess the accuracy of obtained solutions, validation of the computational model is demonstrated against the experimental data of Harris, Ladson & Hill and Mineck & Hartwich (NASA Langley). The comparison is conducted not only for the reference (impermeable) but also for the perforated (permeable) surface NACA 0012 airfoils. Subsequent numerical analysis of SBLI control by moving wall confirms that for the selected velocity ratios, the method is able to improve the shock-upstream boundary layer and counteract flow separation, significantly increasing the airfoil aerodynamic performance.Originality/valueThe moving wall concept as a means of normal shock wave–turbulent boundary layer interaction and shock-induced separation control has been investigated in detail for the first time. The study quantified the necessary operational requirements of such a system and practicable aerodynamic efficiency gains and simultaneously revealed the considerable potential of this promising idea, stimulating a new direction for future investigations regarding SBLI control.

Aeroelastic response of a thin panel excited by a separated shock–boundary layer interaction

Wind-tunnel experiments were conducted to investigate the effects of a separated shock–boundary layer interaction on the aeroelastic behavior of a thin panel in turbulent flow. A suite of traditional sensors and non-contact measurement techniques, including stereo digital image correlation (DIC), were used to characterize the operating conditions and structural response. The impact of the temperature differential between the panel and frame as well as cavity pressure on the panel response were explored. For a majority of the operating conditions, measured time histories of the panel dynamics revealed small-amplitude, forced oscillations about a stationary deformation. These deformations were due to the pressure fluctuations from the turbulent boundary layer and separated shock/boundary layer interaction. However, for a cavity pressure of 83 kPa and temperature differentials of [15, 17.2] K, the response transitioned to bi-stable, cross-well oscillations that were captured for the first time using DIC. The dynamic response was centered about the panel midpoint, coincident with the location of the shock-induced separation on the panel. Using several indirect tests, the intermittent, cross-well oscillations were characterized as potentially chaotic. As the temperature differential continued to decrease, the snap-through motions subsided and the panel response resumed a stationary behavior, albeit at a different deformation shape.

Characteristics of Shock-Induced Boundary-Layer Separation on Nacelles Under Windmilling Diversion Conditions

The boundary layer on the external cowl of an aeroengine nacelle under windmilling diversion conditions is subjected to a notable adverse pressure gradient due to the interaction with a near-normal shock wave. Within the context of computational fluid dynamics (CFD) methods, the correct representation of the characteristics of the boundary layer is a major challenge in capturing the onset of the separation. This is important for the aerodynamic design of the nacelle, as it may assist in the characterization of candidate designs. This work uses experimental data obtained from a quasi-2D rig configuration to provide an assessment of the CFD methods typically used within an industrial context. A range of operating conditions are investigated to assess the sensitivity of the boundary layer to changes in inlet Mach number and mass flow through a notional windmilling engine. Fully turbulent and transitional boundary-layer computations are used to determine the characteristics of the boundary layer and the interaction with the shock on the nacelle cowl. The correlation between the onset of shock-induced boundary-layer separation and the preshock Mach number is assessed, and it was found that the CFD is able to discern the onset of boundary-layer separation.

Evolution of unsteady secondary flow structures near the onset of stall in a tip-critical transonic axial flow compressor stage

The aerodynamic stability of an axial compressor stage depends on the rotor and stator. However, due to the specific design requirements, the evolving flow field and the resulting secondary flow structures are different in both components. This study investigates the evolution of dominant secondary flow structures occurring in the rotor and stator of a tip-critical transonic compressor stage at the near-stall condition using unsteady numerical analysis and validates performance characteristics with the experimental data. The investigation reveals that the presence of rotor tip shock creates a large difference in the pressure gradient across the pressure surface and the suction surface, intensifying tip leakage flow and shock-induced boundary layer separation. The higher incidence angle near the hub leading edge creates a local separation and reattachment zone. The radial pressure gradient causes the low momentum flow from this local separation zone to migrate radially upward. This migrated flow interacts with the tip leakage flow and separated blade boundary layer, eventually creating a colossal recirculation zone and subsequent rotor blockage of around 46%. The increasing streamwise adverse pressure gradient pushes the tip shock upstream, and at the onset of the stall, the flow directly separates from the rotor leading edge avoiding the shock formation. The stator flow field is dominated by the asymmetric hub corner separation induced by the streamwise adverse pressure gradient and the tip corner separation caused by the vortex structures convected downstream from the rotor tip region leading to stator blockage of around 48%. Along with the blade passing frequency, four other dominating frequencies (0.06×BPF, 0.12×BPF, 0.44×BPF, and 0.84×BPF) related to the aerodynamic instabilities are observed at the inception of stall.

Lifetime based pressure and temperature sensitive paint measurement on a civil aircraft wing under conditions near the onset of shock-induced separation

Lifetime-based pressure- and temperature-sensitive paint (PSP and TSP) measurements were conducted in a large-scale industrial transonic wind tunnel to obtain high-quality pressure data for validation of computational fluid dynamics (CFD) simulations. A wind tunnel test was performed using the NASA common research model in the JAXA 2 m × 2 m transonic wind tunnel. The freestream Mach number was varied in the range of 0.70–0.89, and aerodynamic forces and moments were acquired in order to obtain the pitch break condition related to the onset of shock-induced separation. Polymer-based PSP was coated on a main wing of the model, and TSP was used in tandem with PSP to correct the temperature dependence of PSP. The two-gate lifetime-based method was applied to obtain the PSP and TSP emissions. An a-priori/in-situ hybrid calibration was conducted to convert ratio-of-ratios to pressure, and measured pressure distributions were mapped onto a three-dimensional (3D) model grid. The root-mean-square error for the pressure measurements was evaluated by pressure tap data and was approximated to be 0.8 kPa for all Mach numbers tested. The obtained pressure distributions exhibited a significantly high signal-to-noise ratio and were used for comparison with CFD results on a 3D grid. The high-spatial-resolution PSP measurements helped to accurately localize the differences from the CFD simulation results and showed that the prediction of the shock location along the main wing is still a relevant challenge.