Shock Boundary Layer Interaction Research Articles

Numerical studies on shock wave and boundary layer interaction in the high-load turbine rotor cascades

Design of transonic high-load turbine is the essential approach to improve thrust-to-weight ratio of gas turbine engines. The shock wave and boundary layer interaction (SWBLI) in high-load turbine is one of the important unsteady sources and the main source of loss in turbine stages. In order to study mechanism of SWBLI in high-load turbine, rotor blades with loading coefficients range from 1.6 to 2 were designed and Delayed Detached Eddy Simulations (DDES) were carried out. The influences of loading coefficient, exit isentropic Mach number and incidence angle on characteristics of shock wave, SWBLI as well as flow details in the blade passages were compared. Results indicated the shock wave was generated and enhanced as exit isentropic Mach numbers increased leading to the increase of 2.8% in the total pressure loss coefficient. There was not significant difference in the total pressure loss coefficient as loading coefficient increased, however the shock wave was stronger and the separation bubble was longer at the loading coefficient of 1.6. Meanwhile, the influence of strong separation near the leading edge of suction surface induced by variation of incidence angles on the characteristics of the shock wave and the loss was also obvious.

Investigation of Transonic Flow Features in an Axial Flow Compressor Rotor

Numerical investigation of an axial flow rotor is carried out for its performance characterization and aerodynamic behavior during the design and off-design operating conditions. The study focuses on capturing the transonic flow features from the choke point (CP) to the near stall (NS) point in the rotor. This includes the analysis of passage shock structure, its movement in the blade passage with varying back-pressure, shock boundary layer interaction, tip leakage flow structure, and resulting losses. The study is carried out from 60 % to 100 % of the design speed using steady and unsteady RANS simulations. Three turbulence models, namely; SST, k-ε, and Reynolds stress models, are employed. The SST model predicted the closest approximation to the experimental data. The rotor aerodynamic performance is predicted in terms of total pressure ratio, efficiency, and flow contours. Unsteady analysis revealed that the primary and secondary tip leakage vortices, combined with the suction side tip corner separation, are the major instabilities near the stall region.

Replicating transition with modified Spalart–Allmaras model

An algebraic transition model has been developed to preserve the “flow-structure-adaptive” characteristics in “Reynolds-Averaged Navier–Stokes” (RANS) computations for multiple transition mechanisms. The formulation is convenient and plausible in a sense that it relies on the local flow information to trigger transition employing an algebraic intermittency parameter γ rather than a γ-transport equation. The turbulence intensity Tu appearing in γ has been evaluated locally using an empirical relation for the turbulent kinetic energy k, resolving the interaction between local and free-stream turbulence intensities. Splitting γ into low and elevated Tu regimes assists in calibrating the model coefficients as well as minimizing the “trial-and-error” inconsistency, involved in most of the correlation-based transition models for initiating proper simulations. The γ function is directly fed into the production term of a modified Spalart–Allmaras (MSA) turbulence model. The artifact is pivotal to precisely representing the relevant physical aspects of the flow, such as the bypass, natural and separation-induced transitions, and boundary layer (BL) separation and shock BL interactions. Numerical results demonstrate that the MSA transition model maintains decent agreement with other transition and non-transition models available in the literature.

Three-dimensional characteristics of separation vortex generated by crossing shock/boundary layer interaction

Three-dimensional characteristics of separation vortex generated by crossing shock/boundary layer interaction

Aeroacoustic Loading of Impinging Supersonic Boundary-Layer Interaction on Statically Deformed Surfaces

This paper concerns the interaction of an impinging shock wave with a supersonic turbulent boundary layer over several distinct and permanently deformed surfaces, resulting in differences in the shock–boundary-layer interaction and the surface acoustic loading. High-order numerical simulations featuring two-dimensional surface deformations typically encountered in experiments are performed. The deformation amplitudes are up to half the incoming turbulent boundary-layer thickness. The results show that the high-pressure region about the shock impingement is significantly altered and can become narrower or wider depending on the local surface inclination of the deformed panel mode. The surface curvature is found to not significantly affect the separation and reattachment locations of the recirculation bubble. The power spectrum analysis of the pressure fluctuations along the panel’s midspan, where the surface attains the largest deformation amplitude, exhibits a rich and varied response. The pressure power spectrum is amplified in all of the surface deformation modes examined, with the magnitude of the amplification varying in the frequency domain, depending on the location and mode.

A synthetic particle group method for the unsteady interaction between shock wave and boundary layer past a compression ramp

To overcome the lack of resolved turbulent flow structures generated by the improved delayed detached eddy simulation (IDDES) model when simulating the shock wave and boundary layer interaction (SWBLI), a new, nonzonal, embedded synthetic turbulence generation method, named the synthetic particle group method (SPOM), is proposed and validated by simulating the supersonic flow past a flat plate at Mach 2.25. In the attached boundary layer, the SPOs are automatically introduced to trigger the turbulent perturbations. The turbulent boundary layer is successfully resolved and fully developed in a short streamwise range to produce the desired outcome. The effect of grid scale is also studied, and a suitable grid-scale value is recommended. Then, the SPOM is applied to the compression ramp flow at Mach 2.25 with unsteady SWBLI. The results show that the IDDES-SPOM is highly effective, accurate, and able to model scenarios where the original IDDES simulation fails. Furthermore, the effects of the inlet boundary layer thickness on local separation unsteadiness are studied. From the results, the lower frequency motion is attributed to a more fully developed boundary layer.

Study of unsteadiness due to 3-D shock–boundary layer interaction in flow over a square-faced protuberance

Study of unsteadiness due to 3-D shock–boundary layer interaction in flow over a square-faced protuberance

Aerodynamic performance increase over an A320 morphing wing in transonic regime by numerical simulation at high Reynolds number

Purpose This study aims to investigate the morphing concepts able to manipulate the dynamics of the downstream unsteadiness in the separated shear layers and, in the wake, be able to modify the upstream shock–boundary layer interaction (SBLI) around an A320 morphing prototype to control these instabilities, with emphasis to the attenuation or even suppression of the transonic buffet. The modification of the aerodynamic performances according to a large parametric study carried out at Reynolds number of 4.5 × 106, Mach number of 0.78 and various angles of attack in the range of (0, 2.4)° according to two morphing concepts (travelling waves and trailing edge vibration) are discussed, and the final benefits in aerodynamic performance increase are evaluated. Design/methodology/approach This article examines through high fidelity (Hi-Fi) numerical simulation the effects of the trailing edge (TE) actuation and of travelling waves along a specific area of the suction side starting from practically the most downstream position of the shock wave motion according to the buffet and extending up to nearly the TE. The present paper studies through spectral analysis the coherent structures development in the near wake and the comparison of the aerodynamic forces to the non-actuated case. Thus, the physical mechanisms of the morphing leading to the increase of the lift-to-drag ratio and the drag and noise sources reduction are identified. Findings This study investigates the influence of shear-layer and near-wake vortices on the SBLI around an A320 aerofoil and attenuation of the related instabilities thanks to novel morphing: travelling waves generated along the suction side and trailing-edge vibration. A drag reduction of 14% and a lift-to-drag increase in the order of 8% are obtained. The morphing has shown a lift increase in the range of (1.8, 2.5)% for angle of attack of 1.8° and 2.4°, where a significant lift increase of 7.7% is obtained for the angle of incidence of 0° with a drag reduction of 3.66% yielding an aerodynamic efficiency of 11.8%. Originality/value This paper presents results of morphing A320 aerofoil, with a chord of 70cm and subjected to two actuation kinds, original in the state of the art at M = 0.78 and Re = 4.5 million. These Hi-Fi simulations are rather rare; a majority of existing ones concern smaller dimensions. This study showed for the first time a modified buffet mode, displaying periodic high-lift “plateaus” interspersed by shorter lift-decrease intervals. Through trailing-edge vibration, this pattern is modified towards a sinusoidal-like buffet, with a considerable amplitude decrease. Lock-in of buffet frequency to the actuation is obtained, leading to this amplitude reduction and a drastic aerodynamic performance increase.

On the shock wave boundary layer interaction in slightly rarefied gas

The shock wave and boundary layer interaction (SWBLI) plays an important role in the design of hypersonic vehicles. However, discrepancies between the numerical results of high-temperature gas dynamics and experiment data have not been fully addressed. It is believed that the rarefaction effects are important in SWBLI, but the systematic analysis of the temperature-jump boundary conditions and the role of translational/rotational/vibrational heat conductivities are lacking. In this paper, we derive the three-temperature Navier–Stokes–Fourier (NSF) equations from the gas kinetic theory, with special attention paid to the components of heat conductivity. With proper temperature-jump boundary conditions, we simulate the SWBLI in the double cone experiment. Our numerical results show that, when the three heat conductivities are properly recovered, the NSF equations can capture the position and peak value of the surface heat flux, in both low- and high-enthalpy inflow conditions. Moreover, the separation bubble induced by the separated shock and the reattachment point induced by impact between transmitted shock and boundary layer are found to agree with the experimental measurement.

Generalizability of transformer-based deep learning for multidimensional turbulent flow data

Deep learning has been going through rapid advancement and becoming useful in scientific computation, with many opportunities to be applied to various fields, including but not limited to fluid flows and fluid–structure interactions. High-resolution numerical simulations are computationally expensive, while experiments are equally demanding and encompass instrumentation constraints for obtaining flow, acoustics and structural data, particularly at high flow speeds. This paper presents a Transformer-based deep learning method for turbulent flow time series data. Turbulent signals across spatiotemporal and geometrical variations are investigated. The pressure signals are coarsely-grained, and the Transformer creates a fine-grained pressure signal. The training includes data across spatial locations of compliant panels with static deformations arising from the aeroelastic effects of shock-boundary layer interaction. Different training approaches using the Transformer were investigated. Evaluations were carried out using the predicted pressure signal and their power spectra. The Transformer's predicted signals show promising performance. The proposed method is not limited to pressure fluctuations and can be extended to other turbulent or turbulent-like signals.

Physical consistency and invariance in machine learning of turbulent signals

This paper concerns an investigation of the invariance and consistency of deep learning of turbulent pressure fluctuations. The long-short-memory model is employed to predict wall pressure fluctuations across physical regimes featuring turbulence, shock–boundary layer interaction, and separation. The model's sensitivity to the data inputs is examined using different input data sets. Training the deep learning model based on the raw signals from different flow regions leads to large inaccuracies. It is shown that the data must be appropriately pre-processed before training for the deep learning model predictions to become consistent. Removing the mean and using the normalized fluctuating component of the signal, the deep learning predictions not only greatly improved in accuracy but, most importantly, converged and became consistent, provided that the signal sparsity remains within the inertial sub-range of the turbulence energy spectrum cascade. The power spectra of the surface pressure fluctuations reveal that the model provides high accuracy up to a certain frequency for the fully turbulent flow. The deep learning model's consistency is evidenced by being transferable across the various probe positions on the wall despite the significant differences in the turbulent flow properties in the training data set, i.e., signals obtained before, after, and inside the shock–boundary layer interaction regions. The model's prediction consistency and invariance to the turbulent signal training location(s) are promising for applying deep learning models to various turbulent flows.

TURBULENCE MODELING OF SHOCK BOUNDARY LAYER INTERACTION IN HYPERSONIC FLOW

Computational fluid dynamics (CFD) simulations of shockwave turbulent boundary layer interactions using different turbulence modeling including Reynolds-averaged Navier-Stokes (RANS) with different closures (Menter's <i>k</i>-ω-SST, and Spalart-Allmaras) have been investigated. The objective of the study is to assess the capability of current CFD codes, Ansys Fluent and STARCCM, to model hypersonic flow over a large hollow cylinder flare for Mach 6 with Re<sub>∞</sub> = 5.33 × 10<sup>7</sup> The results of heat transfer and pressure data over the surface are presented for two-dimensional, axisymmetric, and three-dimensional RANS models using experimental data of LENS II high Reynolds number shock/Ludweig tunnel at the CUBRC center. Effects of properties of real gas, perfect gas, and different types of mesh have been considered for the assessment.

Control of shock–boundary layer interaction in a transonic compressor cascade via heat transfer at a low Reynolds number

The shock–boundary layer interaction mechanism of a transonic compressor profile at different Reynolds numbers (Re) was investigated by a three-dimensional Reynolds averaged Navier–Stokes simulation. The reason behind the deterioration of the transonic blade performance at low Re was revealed. Based on this, the mechanism of wall heat transfer regulating the interaction between the shock and the boundary layer on the blade surface was investigated to clarify the impact of wall heat transfer on transonic blade performance. The results show that as the Re decreases, the blade performance deteriorates, and a critical Re exists. When the Re exceeds this critical value, the separation structure on the suction side becomes the dominant factor contributing to losses. At this point, wall heat transfer reduces the gas kinematic viscosity near the wall, weakening the stabilizing effect of turbulence dissipation on the separated shear layer. This accelerates the transition process, consequently reducing the size of the separation bubble and decreasing blade losses. Below the critical Re, the passage normal shock approaches the exit of the blade passage and increases in strength, causing severe separation on the pressure side and resulting in a sharp decline in blade performance. At this stage, wall heat transfer can restore the blade’s boosting capability and weaken the passage normal shock. Although this may deteriorate the performance on the suction side to some extent, the degree of separation on the pressure side is significantly reduced, causing the wake region to narrow and improving blade performance. Comparatively, the control effect is more significant at a low Re (Re = 1.5 × 105), with losses reduced by 40.6 % compared to adiabatic wall conditions.

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.

Blade tip leakage flow and heat transfer characteristics over a gas turbine blade at subsonic and transonic exit conditions

The present study numerically investigates the blade tip leakage flow (TLF) over a gas turbine blade cascade at low speed ( M ex = 0.1) and transonic ( M ex = 0.92) blade exit Mach numbers, with and without shroud motion. Various blade tip flow phenomena like flow separation and reattachment, flow choking, shock-boundary layer interaction (SBLI), and tip leakage vortex (TLV) are studied, and their influence on the heat transfer coefficient (HTC) distributions over the blade tip and near tip blade suction surface is investigated. The study has found some new zones of heat transfer due to the interaction of horseshoe vortices (HSV) with the blade tip. In addition to the primary TLV, a secondary TLV is formed, which increases near tip blade suction surface heat transfer. The flow transitions to supersonic speed in the aft portion of the blade tip for the transonic case resulting in complex heat transfer distribution due to SBLI, compared to near-uniform distribution for the low speed case. This supersonic flow results in a smaller flow separation zone near the aft portion of the pressure side tip edge for the transonic case compared to the low speed case. The qualitative nature of HTC near the leading edge of the blade tip resembles for low speed and transonic cases. However, near the trailing edge of the blade tip, the qualitative nature of HTC for the transonic case shows a substantial difference due to supersonic flow compared to the low speed case.

Sweep effects on a canonical shock wave/boundary layer interaction

In this study, we use wall-resolved large-eddy simulations to investigate how the presence of a sweep angle in the upstream flow affects a canonical situation involving shock/boundary layer interaction (SBLI). The chosen configuration is based on the experiments of Bo et al. (2012) where an incident shock impinges on a turbulent boundary layer developing at a freestream Mach number of M=2.7, with a Reynolds number based on the momentum thickness of Reθ=3200. The sweep angle is implemented by adding a spanwise velocity component (crossflow) at the inflow and enforcing periodicity conditions on the side boundaries. This corresponds to an idealized configuration of swept SBLI that is unable to reproduce the conical symmetry found in some real-life applications. Several sweep angles are investigated and compared to a reference unswept case. A viscosity is selected for each case in an effort to keep a constant Reθ across the whole range of sweep angles. The mean flow structure is analysed, showing a strongly skewed flow around the separation zone which increases in size compared to the unswept case. Pressure loads on the wall surface are also higher for swept cases; and the characteristic low-frequencies of the flow separation also slightly increase with significant energy density around a Strouhal number based on the separation length and freestream velocity of StLsep=0.15. Proper orthogonal decomposition (POD) is then applied to obtain more accurate informations on the unsteadiness around the interaction area. This analysis reveals in the swept cases the formation of streamwise-elongated structures related to this new frequency of StLsep=0.15 and moving along the spanwise direction at a convection velocity of about 80% of the transverse freestream velocity.

High-speed shock–boundary-layer interaction over deformed surfaces

This paper concerns the interaction of an oblique shock wave with a supersonic turbulent boundary layer over a thin panel surface, leading to shock–boundary layer interaction and panel buckling. We have performed high-order numerical simulations featuring various static two-dimensional surface deformations typically encountered in experiments. The deformation amplitudes we examined were at least half the height of the incoming turbulent boundary layer thickness. The results show that along the panel midspan, where the maximum deformation amplitude is located, the mean and root mean square pressure are affected by about 10%. Cases for which the pressure at the shock–boundary layer interaction was increased relative to the planar case showed to decrease downstream, and vice versa. Despite the weak response to the mean pressure amplitude, the mean pressure surface contour plots reveal that the streamwise, particularly the spanwise distribution, is affected more noticeably. For example, the surface deformation modes are shown to disrupt the spanwise constant mean pressure, forming higher (or lower) values at either the panel's midspan or edges, depending on the mode. Moreover, the surface curvature leads to a characteristic bending of the spanwise distribution, which can be concave or convex depending on the deformation mode. Analysis of the Reynolds stress anisotropy componentality at different heights from the buckled surface reveals a similar spanwise response of the turbulent velocity fluctuations. The results suggest that the deformation rate plays an important role alongside the deformation amplitude in the turbulent layer and shock–boundary layer interaction.

Temporal characteristics of hypersonic flows over a double wedge with Reynolds number

Laminar hypersonic flows at Mach 7.10 with unit Reynolds numbers of 5.2×104, 1.04×105, and 4.14×105 m−1 over a 30°/55° double-wedge configuration were studied to investigate the spatial–temporal characteristics of the flow in a time-accurate manner. Close comparisons were made between previous kinetic and current continuum approaches to test the validity of the continuum assumption, especially considering the presence of large gradients associated with shock–shock and shock–boundary layer interactions, as well as spanwise instabilities. Previous results from direct simulation Monte Carlo, which inherently predicts rarefied effects such as velocity slip and temperature jumps, were found to be in very close agreement with the current work, even for the lowest Reynolds number. The impact of velocity slip and temperature jumps on flow and surface parameters was investigated, and comparisons were made with a no-slip and constant temperature wall model. The temporal and spatial variation of two- and three-dimensional flows were thoroughly investigated using two-dimensional (2D), three-dimensional (3D) periodic sidewall boundary conditions, and a full 3D configuration consistent with existing experimental data. Close comparisons among the 2D and 3D cases were made. The characteristics of 2D periodic oscillations were reported for the moderate Reynolds number case for the first time. The presence of spanwise instabilities, even at a relatively low free stream pressure of about 100 Pa, establishes that the flow field depends on spanwise effects and is fully 3D. High-fidelity numerical schlieren videos captured strong spanwise oscillations for 3D configurations.

Laser fusion cutting: The missing link between gas dynamics and cut edge topography

In laser cutting, the fundamental role of the gas flow for melt removal and kerf formation is generally accepted. Beyond this vague understanding, however, the underlying physical mechanisms are not yet fully understood. In particular, detailed data concerning the momentum and heat transfer between the gas and melt have seldom been reported. This study addresses the local interactions between the cutting gas and kerf surface (melt film surface) in a fundamental way based on a combined experimental, theoretical, and numerical approach. Typical solid-state laser cut edges are analyzed considering the characteristic surface structures and the basic influences of the gas flow on the global and local melt movement. Here, apparent structures in the micrometer range indicate the effect of vortical gas structures close to the wall. Theoretical investigation of the gas boundary layer is conducted by semiempirical equations and the transfer of basic results from the boundary layer theory. It is shown that the boundary layer is in transition between the laminar and turbulent flow, and local flow separations and shock-boundary layer interactions primarily induce spatially periodic and quasistationary instability modes. An improved numerical model of the cutting gas flow confirms the theoretical results and exhibits good agreement with experimental cut edges, reproducing relevant instability modes and quantifying the local momentum and heat transfer distributions between the gas and melt. With the knowledge gained about the underlying physical mechanisms, promising approaches for improvements of the fusion cutting performance are proposed.

Numerical Simulation of Transonic Compressors with Different Turbulence Models

One of the most commonly used techniques in aerospace engineering is the RANS (Reynolds average Navier–Stokes) approach for calculating the transonic compressor flow field, where the accuracy of the computation is significantly affected by the turbulence model used. In this work, we use SA, SST, k-ɛ, and the PAFV turbulence model developed based on the side-biased mean fluctuations velocity and the mean strain rate tensor to numerically simulate the transonic compressor NASA Rotor 67 to evaluate the accuracy of turbulence modeling in numerical calculations of transonic compressors. The simulation results demonstrate that the four turbulence models are generally superior in the numerical computation of NASA Rotor 67, which essentially satisfies the requirements of the accuracy of engineering calculations; by comparing and analyzing the ability of the four turbulence models to predict the aerodynamic performance of transonic compressors and to capture the details of the flow inside the rotor. The errors of the Rotor 67 clogging flow rate calculated by the SA, SST, k-ɛ, and PAFV turbulence models with the experimental data are 0.9%, 0.8%, 0.7%, and 0.6%, respectively. The errors of the calculated peak efficiencies are 2.2%, 1.6%, 0.9%, and 4.9%. The SA and SST turbulence models were developed for the computational characteristics of the aerospace industry. Their computational stability is better and their outputs for Rotor 67 are comparable. The k-ɛ turbulence model calculates the pressure ratio and efficiency that are closest to the experimental data, but the computation of its details of the flow field near the wall surface is not ideal because the k-ɛ turbulence model cannot accurately capture the flow characteristics of the region of high shear stresses. The PAFV turbulence model has a better prediction of complex phenomena such as rotor internal shock wave location, shock–boundary layer interaction, etc., due to the use of a turbulent velocity scale in vector form, but the calculated rotor efficiency is small.