Shock-induced Boundary Layer Separation Research Articles

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.

Loss Analysis of a Transonic Rotor with a Differential Approach to Entropy Generation

The loss mechanism of transonic axial compressors is a long-standing problem that involves almost all types of entropy generation in fluid flows, such as skin friction, shock waves, shear flows, corner separation, and tip vortices. Primarily, sources need to be identified and quantitative comparisons of their contributions need to be made. For such determination, we propose herein a differential approach to entropy generation, called the “differential approach”. Two case studies are analyzed to determine the applicability of this approach: (1) laminar and turbulent incompressible flows in straight circular ducts and (2) turbulent compressible flows in convergent-and-divergent nozzles with shock waves. The results lead to the following conclusions: (a) Qualitatively, the differential approach works well, and the quantified measure is reliable if it is calculated with quality meshes and a suitable turbulence model. This means that the differential approach can be a good tool for predesign optimization. (b) When shocks occur within flow fields, the shock-induced boundary-layer separation can generate more loss than the shock loss alone. Subsequently, the differential approach is applied to complex flows in the NASA Rotor 67, which is a well-known bench-test transonic rotor. The results show that the differential approach not only determines the local losses and associates the source of losses with the flow structures but also qualitatively compares and identifies the largest contributors. These results provide a theoretical foundation for optimizing rotor design and enhancing stability.

Dual separation control and drag mitigation in high speed flows using viscoelastic materials

Boundary layer separation and friction drag form key delimiting phenomena that subvert the aerial platforms from achieving greater efficiency and accessing wider operation envelope. Both these phenomena are significantly aggravated in supersonic platforms due to the interactions between shock waves with the boundary layer that develops over the vehicle surface and within the engines. The present work demonstrates a new paradigm that leverages the native or programmable material properties of the aerostructures to engender simultaneous reduction in the separation scales and mitigation of skin friction drag. As a first step toward realizing this paradigm, the present work demonstrates, for the first time, the simultaneous skin friction drag mitigation in a Mach 2.5 boundary layer and control of shock induced boundary layer separation, both using viscoelastic implants placed under the flow. It is experimentally demonstrated that the appropriately chosen viscoelastic materials can simultaneously reduce the skin friction coefficient at the measurement location by 11% and mitigate the size of a large-scale separated flow by up to 28%. The reported performance matches the current generation flow effectors in both separation scale and skin friction mitigation. The present study opens a new application space for soft/programmable materials in high speed aerial vehicles.

Numerical Investigation and Optimization of Variable Guide Vanes Adjustment in a Transonic Compressor

In the present work, numerical simulation and optimization was carried out to analyze the mechanism of the variable guide vanes (VGVs) of a transonic compressor. A seven-stage transonic compressor including three-stage VGVs was studied. The VGVs were adjusted individually and jointly under different IGV opening degrees. Changes in performance and shock wave were analyzed, and the coupling effect of the VGV joint adjustment was summarized. Aiming at the maximum efficiency, the joint turning angles were optimized. A novel phenomenon was found wherein the VGV adjustment can affect not only its own performance and that of adjacent downstream blades, but also that of upstream blades. Incidence and performance of upstream blades are improved, but those of the VGV and its adjacent downstream blades are deteriorated. VGV adjustment weakens the shock wave and shock-induced boundary layer separation. The optimal solution for VGV joint adjustment is the combination of the optimal solutions for single VGV adjustments. The joint adjustment optimization improves the efficiency by 0.2–1.93% under different IGV opening degrees.

Nozzle Flow Separation Phenomena and Control for different conditions

A detailed study of separated nozzle flows has been conducted. For a subscale, non-axisymmetric, two-dimensional, convergent divergent nozzle, schlieren flow visualization was acquired along with measurements of force, moment, and pressure as part of an extensive static performance evaluation. Additionally, two-dimensional numerical simulations were performed using the computational fluid dynamics package PAB3D together with algebraic Reynold’s stress modelling and two-equation turbulence closure. This study's experimental findings show that shock-induced boundary layer separation, which was classified into two distinct flow regimes: three-dimensional separation with partial reattachment and entirely detached two-dimensional separation, dominated off design over expanded nozzle flow. The impact of variable shock generation and reflections in various nozzle types on the two primary separation modes, namely Free and Restricted Shock Separation (FSS & RSS), is investigated. The flow separation problem in rocket nozzles has been an unwelcome phenomenon for engineers ever since the birth of the space era. Naturally, the engineers were given the job of bringing things under control. But it proved to be a difficult endeavor; despite the fact that many people were able to explain the physics underlying this occurrence, it is still not completely understood today.

The Bump Suppression Strategy for the Transonic Buffet of the Supercritical Airfoil

Transonic buffet brings challenges to civil aircraft design, including but not limited to the structural fatigue and the handling quality degradation. Therefore, it is necessary to analyze the transonic buffet of the supercritical airfoil and propose the buffet suppression strategy. This paper first employs the delayed detached-eddy simulation method to simulate the unsteady buffet process of the OAT15A supercritical airfoil. Then, the temporal–spatial characteristics of the flowfield and the separation region dynamics are analyzed based on the time-domain analysis method and the dynamic mode decomposition method. After that, different two-dimensional bumps are arranged in the shock motion region to suppress the shock oscillation, and the suppression mechanism and suppression strategy are studied. The results show that the process of the developed transonic buffet is significantly influenced by the merging of the front and the rear separation regions. In case the bump obstructs this confluence phenomenon, the developed transonic buffet can be suppressed. Furthermore, it is also found that a two-dimensional bump with superior buffet performance should have the following features: The bump should be placed after the shock-induced boundary-layer separation point to make the shock-induced boundary-layer separation bubble generate on the bump ramp. Also, the height of the bump should be carefully designed to prevent the front and rear separation regions from merging and to make the front separation region disappear quickly. Finally, this strategy is successfully applied to another supercritical airfoil to confirm the universality of the controlling strategy proposed.

Separation Control for a Transonic Convex Corner Flow Using Ramp-Type Vortex Generators

A flap can be used to control wing camber and as a high-lift device. A convex corner is a simplified model of the upper surface of a flap. At transonic speeds, shock-induced boundary layer separation (SIBLS) occurs at greater freestream Mach numbers and deflection angles. This results in energy losses and a reduction in aerodynamic performance. This study installs ramp-type vortex generators (VGs) upstream of a convex corner, and the effect of the height of the VG on SIBLS is determined. As the height of the VG increases, the magnitude of the mean surface pressure upstream of the corner increases and downstream expansion decreases, which results in a reduction in lift. A reduction in peak surface pressure fluctuations, the separation length, and the frequency of shock oscillation is also determined. For flow control and lift enhancement, micro-VGs are more effective.

Investigation of pressure feedback technique to control ramp based SWBLI

Shock wave boundary layer interaction (SWBLI) and associated boundary layer separation create undesirable phenomena such as enhanced aero-thermodynamic loads, localized turbulent spots, unsteadiness, drag enhancement, pressure recovery losses in high-speed flow applications. It necessitates an investigation on the control of shock-induced boundary layer separation. Therefore, a pressure feedback technique (PFT) to control the boundary layer separation has been studied numerically on a two-dimensional flat plate and ramp configuration. PFT consists of a channel that drains fluid from the separated region and reintroduces it in the core flow under the influence of pressure difference. The effect of various freestream, wall, and geometric conditions on the separation controllability of the PFT is examined. It has been observed that PFT with injection located near the upstream influence point and suction at the ramp foot reduces the length of the separation bubble by 12.15%. Variation of injection location marginally affects the separation reduction capability of PFT, while the performance of PFT is greatly dependent on suction location. The parametric study noted that the optimum suction location, for which the separation size is minimum, shifts downstream with increasing freestream Mach number, wall temperature, and ramp angle. Again the optimum suction location shifts upstream as freestream stagnation temperature to wall temperature ratio increases. Reynolds number is seen to have no impact on the optimum suction location. Moreover, freestream stagnation temperature to wall temperature ratio, which is a governing parameter in SWBLI studies, is also a relevant factor in the mechanism of PFT. Finally, the present limited investigation revealed that the suction end of the PFT positioned between Xsu/L = 1.12 to 1.194 provides a minimum size of shock-induced separation for varying freestream and wall conditions.

Flow structures of strong interaction between an oblique shock wave and a supersonic streamwise vortex

The oblique shock/vortex interaction (OSVI) is numerically investigated based on the large-eddy simulation method. A Mach interaction between separated shock and incident shock can be found when the pressure at the recirculation region reaches a certain level. Based on the idea of spatial–temporal correlation, which considers the three-dimensional steady interaction as a two-dimensional unsteady problem, a qualitative analysis is conducted to explain complicated three-dimensional shock structures. The interaction can be regarded as a combination of the following events: the interaction between circular shock and normal shock, the reflection of shock wave on a subsonic interface, and the interaction between secondary circular shock and other shock structures. Though the original vortex has broken down, a pair of streamwise vortices can be observed in the downstream flow field, the formation of which is associated with the split of the recirculation region. Moreover, the recirculation region is found to act as a solid body, which means that the flow angle along a splitting curve can reflect the splitting speed. Three stages can be identified according to the change process of the flow angle along the splitting curve, which are rapid growth, linear growth, and decrease stages. Inspired by the studies on the shock-induced boundary layer separation, the flow field of the strong OSVI with a regular interaction is modeled to predict the initial flow angle of the splitting point which is the foundation of the study on other stages. The interaction type between separated shock and incident shock can also be judged according to this approach.

Unsteady Aerodynamics Around a Pitching Airfoil with Shock and Shock-Induced Boundary-Layer Separation

This study numerically investigated the characteristics of unsteady aerodynamic forces around a pitching airfoil in the transonic flow regime. The unsteady Reynolds-averaged compressible Navier–Stokes (URANS) equations were solved using a Spalart–Allmaras (SA) turbulence model. The pitching NACA64A010 airfoil at mean angles of attack of 0, 2, 4, and 6 deg with the amplitude of 1 deg and reduced frequency of 0.2 was parametrically simulated. The study found that the trend of the unsteady aerodynamic characteristics around the pitching airfoil, particularly that of the unsteady lift force, significantly changes with the mean-angle-of-attack conditions. The shock-induced boundary-layer separation and reattachment during the pitching oscillation have a crucial role in determining the trend of the unsteady aerodynamic forces, such as the phase-delayed or phase-advanced features for the variation in the angle of attack. The boundary-layer separation during the oscillation, observed at a high mean angle of attack of 6 deg, produces the phase-advanced feature of the lift force. Because the phase-advanced feature is associated with a positive damping effect, the result indicates that the boundary-layer separation may serve as a stabilizing effect on the pitching airfoil motion. The boundary-layer reattachment from separation during the airfoil oscillation, observed at an intermediate mean angle of attack of 4 deg, generates an unusual shock wave movement, resulting in a unique lift loop for the variation in the angle of attack. The study also demonstrated a good prediction accuracy of the URANS with the SA model for the unsteady aerodynamics around the pitching airfoil with the shock and shock-induced boundary-layer separation.

Quasi-wall-resolved large eddy simulation of transitional flow in a transonic compressor rotor

The large eddy simulation (LES) method is a high-resolution and promising approach for the accurate turbulence analysis of axial-flow compressors. However, the validity of LES method in the analysis of the laminar-turbulent transition in the high-speed compressor is unknown or not fully exploited, and the flow physics associated with the interaction between the transition and the shock-induced boundary layer separation remains unclear. In the present work, a concept of quasi-wall-resolved LES (QWRLES) method is developed, in which the near-wall grid resolution lies between that of wall-resolved LES (WRLES) and wall-modeled LES (WMLES) but the height of the first layer of near-wall grid cell is within 10 in wall units to permit a direct calculation of wall-shear stress via the Newtonian constitutive relation. The prediction accuracy of QWRLES is first verified in the channel-flow and transonic bump-flow problems. Then, the simulation of the transonic compressor NASA Rotor 37 is carried out to validate the QWRLES method and to explore the transition mechanism. Reasonable agreement between the calculated and measured aerodynamic performance is achieved. The simulation results show that the transition occurs in a bypass mode on the blade suction surface and near the shroud- and hub-endwalls. The shock-boundary layer interaction, the strong shear in the tip-leakage region and the pitchwise shear of the rotating hub are responsible for the transition in the above three regions, respectively. The simulation results show high resolution of the LES method for the transitional and turbulent flows in axial-flow compressors.

Stall Behavior in an Ultrahigh-Pressure-Ratio Centrifugal Compressor: Backward-Traveling Rotating Stall

AbstractThis paper describes the stall mechanism in an ultrahigh-pressure-ratio centrifugal compressor, composed of a double-splitter impeller, radial diffuser, and axial diffuser. A model comprising all impeller and diffuser blade passages is used to conduct unsteady simulations that trace the onset of instability in the compressor. Backward-traveling rotating stall waves appear at the inlet of the radial diffuser when the compressor is throttled. Six stall cells propagate circumferentially at approximately 0.7% of the impeller rotation speed. The detached shock of the radial diffuser leading edge and the number of stall cells determine the direction of stall propagation, which is opposite to the impeller rotation direction. Dynamic mode decomposition is applied to instantaneous flow fields to extract the flow structure related to the stall mode. This shows that intensive pressure fluctuations concentrate in the diffuser throat as a result of changes in the detached shock intensity. The diffuser passage stall and stall recovery are accompanied by changes in incidence angle and shock wave intensity. When the diffuser passage stalls, the shock-induced boundary–layer separation region near the diffuser vane suction surface gradually expands, increasing the incidence angle and decreasing the shock intensity. The shock is pushed from the diffuser's throat toward the diffuser leading edge. When the diffuser passage recovers from the stall, the shock wave gradually returns to the diffuser throat, with the incidence angle decreasing and the shock intensity increasing. Once the shock intensity reaches its maximum, the diffuser passage suffers severe shock-induced boundary–layer separation and stalls again.

Micro-Vortex Generators on Transonic Convex-Corner Flow

A convex corner models the upper surface of a deflected flap and shock-induced boundary layer separation occurs at transonic speeds. This study uses micro-vortex generators (MVGs) for flow control. An array of MVGs (counter-rotating vane type, ramp type and co-rotating vane type) with a height of 20% of the thickness of the incoming boundary layer is installed upstream of a convex corner. The surface pressure distributions are similar regardless of the presence of MVGs. They show mild upstream expansion, a strong favorable pressure gradient near the corner’s apex and downstream compression. A corrugated surface oil flow pattern is observed in the presence of MVGs and there is an onset of compression moving downstream. The counter-rotating vane type MVGs produce a greater reduction in peak pressure fluctuations and the ramp type decreases the separation length. The presence of MVGs stabilizes the shock and shock oscillation is damped.

The Effect of Vortex Generators on Shock-Induced Boundary Layer Separation in a Transonic Convex-Corner Flow

Deflected control surfaces can be used as variable camber control in different flight conditions, and a convex corner resembles a simplified configuration for the upper surface. This experimental study determines the presence of passive vortex generators, VGs (counter-rotating vane type), on shock-induced boundary layer separation for transonic convex-corner flow. The mean surface pressure distributions in the presence of VGs for h/δ = 0.2 and 0.5 are similar to those for no flow control. If h/δ = 1.0 and 1.5, there is an increase in the amplitude of the mean surface pressure upstream of the corner’s apex, which corresponds to greater device drag and less downstream expansion. There is a decrease in peak pressure fluctuations as the value of h/δ increases, because there is a decrease in separation length and the frequency of shock oscillation. The effectiveness of VGs also depends on the freestream Mach number. For M = 0.89, there is an extension in the low-pressure region downstream of a convex corner, because there is greater convection and induced streamwise vorticity. VGs with h/δ ≤ 0.5 are preferred if deflected control surfaces are used to produce lift.

Global Visualization of Compressible Swept Convex-Corner Flow Using Pressure-Sensitive Paint

This study used pressure-sensitive paint (PSP) and determined the surface pressure distributions for a compressible swept convex-corner flow. The freestream Mach numbers were 0.64 and 0.83. The convex-corner angle and swept angle were, respectively, 10–17° and 5–15°. Expansion and compression near the corner apex were clearly visualized. For the test case of shock-induced boundary layer separation, there were greater spanwise pressure gradient and curved shocks. The acquired PSP data agree with the experimental data measured using the Kulite pressure transducers for a subsonic expansion flow. For a transonic expansion flow, the discrepancy was significant. The assumption of a constant recovery factor is not valid in the separation region, and temperature correction for PSP measurements is required.

A discontinuous Galerkin method for wall-modeled large-eddy simulations

We develop an augmented discontinuous Galerkin method for wall-modeled large-eddy simulations. This method is motivated by the enrichment method that has been formulated for finite-element method, by statistically augmenting the variational solution of the near-wall element with a simple wall function to model the effects of the unresolved momentum-carrying near-wall eddies on the resolved scales. An explicit model is employed to model the subgrid-scale stresses. It is shown that this approach improves the solution accuracy on significantly underresolved grids and reduces spurious oscillations that occur in the presence of strong wall gradients on coarse grids with high-order DG methods. The proposed method is applied in low-Mach-number plane channel flows over a range of Reynolds numbers to examine effects of mesh resolution, polynomial order, solution-augmentation, and the SGS closure on the model performance. The results show that the proposed approach reproduces the log-law profile for different flow conditions, even at extreme Reynolds numbers, and yields improved predictions with increasing polynomial order. In addition, the capability of the present DG-based WMLES framework for computations of complex flows is demonstrated in LES of a shock-induced boundary-layer separation for the transonic bump experiment by Bachalo & Johnson (AIAA J., 24, 1986).

Control of Shock-Induced Separation of a Turbulent Boundary Layer Using Air-Jet Vortex Generators

An experimental analysis was carried out to study the control effectiveness and turbulence behavior of an array of steady air-jet vortex generators on a shock-induced boundary-layer separation. Control was applied to a 24° compression-ramp interaction at Mach 2.52 and . A parametric study varying the jet spacing revealed that an adequate interaction between the vortices is essential for effective control, and the most favorable control effect was achieved with a jet spacing of . This case was analyzed in more detail using particle image velocimetry. The results reveal a strong corrugation of the separation zone due to the injection of the jets with a substantial reduction in its area along the entire spanwise extent of the air-jet array, compared with the uncontrolled case. Measurements of velocity fluctuations along the interaction region also show that the jet injection results in a reduction in turbulent intensity after reattachment, contributing to an overall reduction in amplification factor across the interaction region. The expected increase in turbulent mixing, which is essentially responsible for the control effect, was confirmed in the region in-between jet injection and separation shock. Farther away from the surface, the flow is not affected, which reduces potential drag penalties.

Experimental Investigation of Microramp Control of an Axisymmetric Shock/Boundary-Layer Interaction

Mitigation of shock-induced boundary-layer separation using vortex generators (VGs) is an emerging flow control approach in planar inlet geometries. In the present work, an experimental campaign is performed to determine the effectiveness of microramp VGs at shock-induced separation control in an axisymmetric configuration. The baseline flowfield consists of a shock-induced separation generated by a 20 deg axisymmetric ramp mounted within a half-cylindrical duct at and . The near-wall VG flowfield and pressure field are first characterized to ensure consistency with planar microramp configurations. The effects of streamwise placement and single/multiple VGs on the shock-/boundary-layer interaction are subsequently investigated. The results indicate that, although the devices are effective at reducing separation immediately downstream, they cause an increased separation shock strength and interaction pressure rise, which enlarge separation in the uncontrolled areas of the flow.

Humidity effects on the aerodynamic performance of a transonic compressor cascade

Humid air degrades gas turbine aerodynamic performance. However, little is known about humidity effects on the compressor flow field. Blade loading distribution, deviation, and loss coefficient have been measured in a transonic compressor cascade for relative humidity (RH) values of 20, 32, 42, 45, 49 and 53% at Mach number of 1.11 and incidence of +5°. In addition, numerical studies have been conducted for the same conditions by incorporating the Classical Nucleation Theory (CNT) model and Hertz-Knudsen droplet growth model into ANSYS FLUENT. As the relative humidity increases, due to condensation, the pressure of upstream of the shock on the suction surface is increased; the pressure on the pressure surface is decreased; and shock is pushed downstream, reducing shock strength. However, due to the negligible shock-induced boundary layer separation, humidity has little effect on deviation. Loss coefficient is increased as the relative humidity increases because, even though the shock loss is reduced, heat addition to the flow from condensation decreases total pressure.