Shock Wave Boundary Layer Interaction Research Articles

Wall pressure fluctuations in wall heated and cooled shock wave and turbulent boundary layer interactions

Wall pressure fluctuations in shock wave and turbulent boundary layer interactions on heated and cooled walls are investigated by performing direct numerical simulations. The objectives of this study are to identify flow dynamics causing low- and high-frequency wall pressure fluctuations and evaluate wall heating and cooling effects on the wall pressure fluctuations. The wall pressure spectra highlight two characteristic low-frequency spectral peaks at the separation point and below the expansion waves. Conditional analysis reveals that the two spectral peaks originate from the oscillations of the reflected shock and expansion waves, coupled with the expansion and contraction motions of the separated turbulent shear layer. Regarding wall temperature effects, wall cooling enhances the magnitude of the wall pressure fluctuations regardless of the low- and high-frequency components. Furthermore, influences of the expansion waves are also enhanced by wall cooling, which increases the wall pressure fluctuations locally after the separation.

Control of Cowl Shock/Boundary Layer Interaction in Supersonic Inlet Based on Dynamic Vortex Generator

A shock wave/boundary layer interaction (SWBLI) is a common phenomenon in supersonic inlet flow, which can significantly degrade the aerodynamic performance of the inlet by inducing boundary layer separation. To address this issue, in this paper, we propose the use of a dynamic vortex generator to control the SWBLI in a typical supersonic inlet. The unsteady simulation method based on dynamic grid technology was employed to verify the effectiveness of the proposed method of control and investigate its mechanism. The results showed that, in a duct of finite width at the inlet, the SWBLI generated complex three-dimensional (3D) flow structures with remarkable swirling properties. At the same time, vortex pairs were generated close to the side wall as a result of its presence, and this led to the intensification of transverse flow and, in turn, the formation of a complex 3D structure of the flow of the separation bubble. The dynamic vortex generator induced oscillations of variable intensity in the vortex system in the supersonic boundary layer that enhanced the mixing between the boundary layer flow and the mainstream. Meanwhile, the unique effects of “extrusion” and “suction” in the oscillation process continued to charge the airflow, and the distribution of velocity in the boundary layer significantly improved. As the oscillation frequency of the vortex generator increased, its charging effect on low-velocity flow in the boundary layer increased, and its control effect on the flow field of the SWBLI became more pronounced. The proposed method of control reduced the length of the separation bubble by 31.76% and increased the total pressure recovery coefficient at the inlet by 6.4% compared to the values in the absence of control.

Modelling and shock control for a V-shaped blunt leading edge

Hypersonic flow on a V-shaped blunt leading edge (VSBLE) at Mach 12 is numerically investigated. A series of self-induced shock–shock interactions and shock wave/boundary layer interactions (SWBLIs) cause extremely high heat flux peaks on the crotch of the VSBLE. Based on these shock structures, a simplified model that divides the flow field into inviscid and viscous areas is constructed. This model can quickly solve the shock interactions and predict the heating/pressure peaks with high accuracy. Furthermore, a shock control bump (SCB) is placed on the SWBLI region to split the strong incident shock into a weaker multi-wave system. The mechanism study of the SCB shows that the inviscid effect significantly reduces the heating/pressure peaks, and the viscous effect suppresses the SWBLI-induced separation. Finally, a VSBLE with SCBs is numerically investigated. The heat flux peak is reduced by 66 % compared to that without the SCBs. The robustness of the SCB under various working conditions is also evaluated. This paper provides an idea for the simplified solution of complex shock interactions and extends the application of the SCB as a thermal protection device in hypersonic flow for the first time.

Supersonic Flow Separation in Front of a Forward-Facing Step

A detailed experimental study has been done to understand the Shock Wave-Boundary Layer Interaction (SWBLI) over a Forward-Facing Step (FFS) of height (h) in a Mach 2.5 flow using Particle Image Velocimetry (PIV). In addition to PIV, unsteady wall pressure measurements, high-speed schlieren and surface oil flow visualization have been studied. Flow separation using surface oil flow visualization was observed around 4.1 step heights upstream of the step face. Shock motions extracted from the schlieren image sequence shows a peak frequency of approximately 1000 Hz, which corresponds to the dominant (low) frequency at which the shock oscillates; this being two orders in magnitude smaller than the characteristics frequency of the incoming boundary layer. Also, spectra of the unsteady wall pressure data shows that the shock oscillates with a peak frequency of approximately 1000 Hz. The effect of shock location (xs) from instantaneous velocity PIV fields in the cross-stream plane due to an upstream influence parameter (line-averaged boundary layer streamwise velocity, ul) and a downstream parameter (separation bubble or recirculation region area, A) is studied. Instantaneous cross-stream PIV measurements show significant variations of the separation shock location, incoming boundary layer velocity fluctuations and recirculation region. Besides, conditionally averaged velocity field based on incoming boundary layer streamwise velocity fluctuations and separation bubble area in the cross-stream plane shows that the shock location (xs) is correlated to the size of the separation bubble and not to boundary layer velocity fluctuations.

Selected results of the hypersonic flight experiment STORT

The three-stage rocket configuration of the hypersonic experiment STORT with several scientific payloads concerning hypersonic technologies was launched successfully from the Andøya Space launch site in northern Norway on June 26, 2022. The third stage performed a suppressed trajectory to increase integral heat load on the payload structures. The vehicle traveled at speeds above Mach 8 at altitudes between 30 km and 38 km for more than 60 s. The nose and forebody section of the payload was made of CMC structures. Three canards with a CMC thermal protection were equipped with thermal management experiments to verify the thermal efficiency of these methods. All engineering science experiments like aerothermal heating of the nose and forebody, CMC material response at temperatures above 2200 K, hypersonic thermal management, shock wave boundary layer interaction, CFRP module with cork coating for high temperature applications, high temperature fin leading edge and radiometer sensors provided unique flight data. A reduced in-flight 3 DoF trajectory simulation running on the flight computer determined the ignition time of the third stage which allowed improved experiment conditions in terms of Mach number and apogee altitude even in the presence of external perturbations.

Numerical study of the effect of micro vortices on chaotic flutter

The loading generated by shock wave boundary layer interaction (SBLI) plays a crucial role in the structural design of a high-speed vehicle. They can cause flow separation near the shock impingement, resulting in the onset of chaotic, self-sustained flutter leading to structural failure. This work investigates how a micro-ramp installed upstream of a flexible panel affects its flutter behavior. For this purpose, two-way coupled laminar fluid–structure interaction simulations are performed for two different oblique shock strengths impinging over a flexible panel. A comprehensive study with a micro-ramp, as high as half the boundary layer thickness, is conducted for a wide range of non-dimensional dynamic pressures expected for a supersonic flight. The study provides analysis methodology to characterize the nature of panel oscillations, and the influence of micro-ramp on fluid and structural unsteadiness is evaluated. The results indicate that the chaotic flutter is initiated because of the non-linear interaction between the high-frequency fluid and low-frequency structural unsteadiness. Micro vortices delay the onset of the chaotic flutter by lowering the fluid frequency, thereby synchronizing fluid and structure unsteadiness. The micro vortices also decreased flutter frequency for lower shock strength; however, MVG was somewhat ineffective for some higher shock strength. For the latter taller micro-ramps that can introduce stronger, low-frequency vortices are recommended.

Effects of favourable downstream pressure gradients on separated shock-wave/boundary-layer interactions

Shock-wave/boundary-layer interactions (SWBLI) strongly influence the aerodynamic behaviour of many aerospace, propulsion and transportation systems. They have therefore been studied intensively — with different methods, configurations, and test facilities. Regions of favourable pressure gradient are often placed downstream of the SWBLI region., e.g. to prevent wind-tunnel blockage. To be able to better understand their effect and compare cases from varying sources, a joint experimental–numerical study of a supersonic compression/expansion corner flow is carried out. A high-resolution long-time numerical simulation complements particle-image velocimetry measurements, both using the same configuration and conditions. Derived setups without the downstream expansion corner are also explored; both for the same compression-corner angle and for a ramp angle inducing a separation bubble of the same size as for the original compression/expansion configuration. Results demonstrate that the topology and very low-frequency dynamics of the separation bubble developing over the compression corner are altered by the downstream expansion corner. Differences are mainly found for the downstream motion of the shock/shrinking of the separated region and extend downstream of the reattachment point up to the expansion corner.

Experimental investigation on effects of herringbone riblets on shock wave/boundary layer interactions of a compression ramp at Mach 3

In this paper, the effects of herringbone riblets (HRs) on a turbulent boundary layer and the shock wave/boundary layer interactions (SWBLIs) are experimentally investigated at the Mach number of 3. An array of three varied lengths of HRs strips are applied upstream of the separation zone of SWBLIs on a compression ramp model. High-speed schlieren, oil-flow visualization, and the planar laser scattering technique are used to examine the shock pattern and boundary layer developing over the surface of the model. The snapshot proper orthogonal decomposition technique and the Fast Fourier Transform method are applied to study the impact of HRs on the interaction between the shock wave and the boundary layer. The experiments provide convincing evidence that HRs make the separation line wavy and shrink the separation zone by about −39.54% for the longest HRs in the present experiments. Furthermore, it is also revealed that these microscale HRs induce large-scale streamwise vortical structures within the boundary layer as found in incompressible flows. It is believed that these vortices promote momentum transfer within the boundary layer hence providing the dominant mechanism for suppressing flow separation.

Improvement of turbulence model for predicting shock-wave–boundary-layer interaction flows by reconstructing Reynolds stress discrepancies based on field inversion and machine learning

Reynolds-averaged Navier–Stokes (RANS) models have been the mainstay of engineering applications in recent years, and this trend will likely persist in the coming decades. However, the ability of RANS methods to predict complex flows, such as shock-wave/boundary-layer interaction (SWBLI) flows, is inadequate. In the absence of a breakthrough in traditional turbulence modeling, data-driven modeling has emerged as a new paradigm in turbulence research. In this study, a field-inversion and machine-learning framework based on the regularized ensemble Kalman filter (REnKF) was implemented to enhance the predictive ability of the Menter shear-stress transport (SST) model for SWBLI flows. This approach directly modifies the norm and eigenvalues of the Reynolds stress tensor obtained by RANS methods to overcome the limitations of the Boussinesq hypothesis. The spatial distributions of Reynolds stress discrepancies are initially obtained from experimental data using the REnKF method and a parameterization method based on geometric transformation for improved prediction of the training case. Then, a mapping function from local flow variables to discrepancy fields is constructed by selecting appropriate input features and combining them with a random forest algorithm. Our results verify the effectiveness of the framework for SWBLI flows of varying strengths and types. The results also demonstrate that the SST model's simulation ability for SWBLI flows has been significantly improved, providing more accurate predictions of separation and reattachment, considerably lowering the prediction errors for relevant flow variables, and verifying the generalization ability of the proposed framework.

Numerical Study of Shock Wave Boundary Layer Interaction Based on an Anisotropic Turbulence Model

This paper presents numerical simulations of incident oblique shock wave impinging on a flat plate with a turbulent boundary layer using a nonlinear anisotropic turbulence model. Three cases with different incident shock angles under the same free-stream Mach number are studied, including attachment flow, incipient separation flow and outstanding separation flows. Convection terms and diffusion terms are calculated using the third-order Roe’s flux difference splitting scheme and the second-order central difference scheme in the code, respectively. The Runge-Kutta time marching method is employed to solve the governing equations for steady state solutions. Comparisons between the simulated results and experimental data are carried out, which include density contours, pressure distribution, skin friction coefficients distribution and streamline vectors. To capture the essential behavior of this complex anisotropic turbulence separation flow caused by the shock-wave-boundary-layer-interaction (SWBLI), an anisotropic eddy viscosity tensor and a nonlinear Reynolds stress-strain constitutive relationship are used in this model. The anisotropic eddy viscosity tensor is composed of turbulent fluctuation velocity vector and characteristic length vector which are obtained by statistical partial average scheme. With proper modeling of anisotropic eddy viscosity tensor, the numerical results are in good agreement with the experimental data. It is worthy to point it out that the computation accuracy decreases with increasing interaction intensity.

Study on Optimization Design of Airfoil Transonic Buffet with Reinforcement Learning Method

Transonic buffet is a phenomenon of large self-excited shock oscillations caused by shock wave-boundary layer interaction, which is one of the common flow instability problems in aeronautical engineering. This phenomenon involves unsteady flow, which makes optimal design more difficult. In this paper, aerodynamic shape optimization design is combined with reinforcement learning to address the problem of transonic buffet. Using the deep deterministic policy gradient (DDPG) algorithm, a reinforcement learning-based design framework for airfoil shape optimization was constructed to achieve effective suppression of transonic buffet. The aerodynamic characteristics of the airfoil were calculated by the computational fluid dynamics (CFD) method. After optimization, the buffet onset angles of attack of the airfoils NACA0012 and RAE2822 were improved by 2° and 1.2° respectively, and the lift-drag ratios improved by 83.5% and 30% respectively. Summarizing and verifying the optimization results, three general conclusions can be drawn to improve the buffet performance: (1) narrowing of the leading edge of the airfoil; (2) situating the maximum thickness position at approximately 0.4 times the chord length; (3) increasing the thickness of the trailing edge within a certain range. This paper established a reinforcement learning-based unsteady optimal design method that enables the optimization of unsteady problems, including buffet.

Effects of expansion waves on incident shock-wave/boundary-layer interactions in hypersonic flows

The effects of expansion waves on incident shock-wave/boundary-layer interactions (SWBLIs) at a Mach number of 4.96 are experimentally studied. The flow characteristics of the interaction zone under various intensities and positions of the incident shock wave and expansion waves are quantitatively analyzed. The expansion waves weaken the intensity of the shock waves encountered, thereby weakening the intensity of the SWBLIs. With an increasing distance between the expansion waves and the interaction zone, the total wall pressure jump and the interference length show a linear growth trend. However, the expansion waves do not affect the initial pressure jump of the separation, which is consistent with free-interaction theory. Finally, the scaling model proposed by Souverein et al. [“A scaling analysis for turbulent shock-wave/boundary-layer interactions,” J. Fluid Mech. 714, 505 (2013)] is simply modified using the measured value of the pressure jump. This correction provides a better approximate result for SWBLIs under the impact of expansion waves.

Vibration of a thin panel exposed to ramp-induced shock-boundary layer interaction at Mach 2

Flight vehicles that operate in the supersonic regime can be subject to adverse fluid–structure interactions due to their lightweight design. The presence of geometric obstructions such as control surfaces or fins can induce shocks that can interact with the boundary layer, leading to flow separation. This work investigates experimentally the interaction between a compliant panel in a Mach 2 flow under a compression ramp-induced shock-wave/boundary-layer interaction (SBLI). Thin brass panels of different thickness are investigated in a wind tunnel. Tests are performed both with and without a 20° compression ramp installed. This direct comparison allows characterization of the effect of the SBLI on the system dynamics. High-speed stereoscopic digital image correlation (DIC) and fast-response pressure sensitive paint (PSP) are used to obtain simultaneous measurements of full field deformation and surface pressure of the panels. The panel vibration is dominated by the first bending mode. Despite the forcing of the separation shock foot, the presence of the SBLI does not significantly modify the operational deflection shape, frequency, and amplitude of the dominant vibration mode, beyond what is observed for the no-SBLI case. On the other hand, analysis of the shock foot motion shows that the shock primarily oscillates at the first natural frequency of the panel. This leads to the conclusion that the shock foot oscillation is driven by the panel vibration in a one-way coupling mechanism. The SBLI does modify the higher modes, which is likely due to localized forcing by the separation shock foot. Full-field surface pressure predictions are made using first order piston theory. Results show that the fluid–structure interaction is dominated by the large region of attached flow upstream of the shock foot.

Unsteady interaction mechanism of transverse stage separation in hypersonic flow for a two-stage-to-orbit vehicle

Due to the complex aerodynamic interaction, the safe separation of two stages is one of the challenges for the successful launch of a two-stage-to-orbit (TSTO) vehicle. The unsteady hypersonic flow past the parallel-staged TSTO model during stage separation at Ma = 6.7 and Re = 8.86 × 105 m−1 is numerically studied using laminar flow simulation. The TSTO model consisted of a waverider and a spaceplane as booster and orbiter, respectively. The effect of the center of gravity (CoG) of the orbiter on the unsteady aerodynamic interference during stage separation of TSTO is analyzed in detail with 0.65 ≤ lCoG/lo ≤ 0.80. In addition, the aerodynamic characteristics, dynamic behaviors, and unsteady wall pressure variation are compared in different cases. The results show that the CoG regime is limited to 5% of the orbiter length for absolutely safe separation, i.e., 0.70 < lCoG/lo < 0.75. As for the unsuccessful separation, the orbiter tends to fly nose-down if lCoG/lo ≤ 0.70 while tending to pitch or somersault when lCoG/lo = 0.80. Furthermore, the pitching moment of the orbiter, which is influenced by the interstage shock wave–boundary layer interaction and shock–shock interaction, dominates the separation safety, and the specific flow mechanisms concerning the separation behavior associated with aerodynamic interference in different cases are analyzed in detail.

Numerical Investigation of Asymmetric Mach 2.5 Turbulent Shock Wave Boundary Layer Interaction

Supersonic shock wave boundary layer interactions are common to inlet flows of supersonic and hypersonic vehicles. This paper reports on wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 turbulent shock wave boundary layer interaction experiment at the NASA Glenn Research Center. The boundary layer upstream of the interaction was nominally axisymmetric and two-dimensional. A conical centerbody with a 16 deg half-angle and a maximum radius of 0.147D of the test section diameter was employed to generate a conical shock wave, where D is the test section diameter. Asymmetric (swept) interactions were obtained by displacing the shock generator away from the test section centerline. The present simulation is for a shock generator displacement of D/6. Results from the asymmetric simulation are compared with results from an earlier simulation of a corresponding axisymmetric interaction. The experimental Reynolds number based on test section diameter was ReD=4×106. For the simulations, the Reynolds number was lowered to ReD=4×105 to keep the computational expense of the simulations within limits. Compared to the axisymmetric interaction, the streamwise extent of the separation varies considerably in the azimuthal direction for the asymmetric interaction. The separation is strongest at the azimuthal location that is closest to the shock generator. The streamwise extent of the separated flow regions is noticeably reduced and substantial crossflow is observed between the locations that are closest and farthest from the shock generator. A Fourier analysis of the unsteady flow data indicates low-frequency content for the separated region that is closest to the shock generator. Away from this region, with increasing sweep angle and cross-flow, the low-frequency content is diminished. A proper orthogonal decomposition captures spanwise coherent structures for the more two-dimensional parts of the interaction.

Shock impingement on a transitional hypersonic high-enthalpy boundary layer

Shock-wave boundary layer interactions are of utmost importance for both aeronautical and aerospace applications. The dynamics of high-speed compressible turbulent boundary layers becomes tightly coupled with strong gradients of thermodynamic properties, leading to an increase of thermomechanical loads. In the contextual presence of hypersonic and high-enthalpy regimes, thermochemical nonequilibrium effects must be considered as well. This work is a first step toward the fundamental understanding of this configuration.

Adaptive control mechanism of shock wave/boundary layer interaction induced by the secondary recirculation jet with outlet backpressure

The increased and unstable flow field backpressure will cause problems such as the non-starting of the inlet tract, and the widespread shock wave/boundary layer interaction (SWBLI) phenomena in the supersonic flow field exacerbates these problems. Hence, a powerful flow control system is required. In this paper, backpressure is introduced at the flow field outlet, and the effect of different backpressure ratios on the flow field is explored. An adaptive control scheme is also developed by using the optimized secondary flow recirculation configuration. The three-dimensional implicit Reynolds Averaged Navier–Stokes equations are utilized for numerical simulation of the flow field. The results show that the adaptive control of the secondary recirculation jet has a positive control effect on the SWBLI of the flow field when backpressure is applied. Moreover, the adaptive control mechanism under the backpressure condition is analyzed, which is applicable to different backpressure flow fields with Mach numbers between 2.5 and 3.5.

Numerical Investigation of Mach 2.5 Axisymmetric Turbulent Shock Wave Boundary Layer Interactions

Shock wave boundary layer interactions are common to both supersonic and hypersonic inlet flows. Wall-resolved implicit large-eddy simulations of a canonical Mach 2.5 axisymmetric shock wave boundary layer interaction experiment at Glenn Research Center were carried out. A conical shock wave was generated with axisymmetric centerbodies with 16 deg half-angle cone. The centerbody radii were 9.2% and 14.7% of the test section diameter. The conical shock wave interacted with the turbulent boundary layer on the inside of the cylindrical test section. The experimental Reynolds number based on diameter was six million. For the simulations, the Reynolds number was reduced by a factor of 10 to lower the computational expense. The turbulent boundary layer separates for both centerbody radii and the separation is stronger for the larger centerbody radius. Frequency spectra of the spanwise-averaged wall-pressure coefficient reveal low-frequency content at Strouhal numbers based on separation length between 0.02 and 0.05 in the vicinity of the separation shock and mid-frequency content between 0.1 and 0.2 downstream of separation. A proper orthogonal decomposition captures spanwise coherent structures with a Strouhal number of 0.03–0.04 over the interaction region and streamwise coherent structures inside and downstream of the interaction with a Strouhal number of 0.1–0.4.

Boundary layer control for low Reynolds number fan rig testing

Ultra-high bypass ratio turbofans offer significant reductions in fuel and pollution due to their higher propulsive efficiency. Short intakes might lead to a stronger fan-intake interaction, which creates uncertainty in stability at off-design conditions. Due to the prohibitive cost of full-scale experimental testing, subscale testing in wind tunnels is used to understand this behaviour. The low Reynolds number of subscale models results in unrepresentative laminar shock-boundary layer interactions. The boundary layer state thus needs to be conditioned to better represent full-scale transonic fans. This paper proposes the use of an inexpensive and robust flow control method for the suction side of a fan blade. Design guidelines are given for the location and height of the discrete roughness elements used to control the boundary layer state. This paper also presents a rapid experimental validation methodology to ensure and de-risk the application of the boundary layer trip to 3D rig blades. The experimental methodology is applied to a generic aerofoil representative of a fan tip section. The experimental method proves that it is possible to reproduce boundary layers and pressure distributions of a full-scale fan blade on a 1/10 subscale model. The results obtained confirm that the boundary layer trip method successfully promotes transition at the location representative of full-scale blades, avoiding unrepresentative laminar shock wave boundary layer interactions. This highlights the importance of conditioning boundary layers in low Reynolds number fan rig testing.

On low-frequency unsteadiness in swept shock wave–boundary layer interactions

We derive a scaling law for the characteristic frequencies of wall pressure fluctuations in swept shock wave/turbulent boundary layer interactions in the presence of cylindrical symmetry, based on analysis of a direct numerical simulations database. Direct numerical simulations in large domains show evidence of spanwise rippling of the separation line, with typical wavelength proportional to separation bubble size. Pressure disturbances around the separation line are shown to be convected at a phase speed proportional to the cross-flow velocity. This information is leveraged to derive a simple model for low-frequency unsteadiness, which extends previous two-dimensional models (Piponniau et al., J. Fluid Mech., vol. 629, 2009, pp. 87–108), and which correctly predicts growth of the typical frequency with the sweep angle. Inferences regarding the typical frequencies in more general swept shock wave/turbulent boundary layer interactions are also discussed.