Shock Boundary Layer Interaction Research Articles

Spatial and temporal alterations due to vortex generators in a flare induced shock–boundary layer interaction

This paper describes the modifications instigated in a flare induced shock–boundary layer interaction flowfield by two vortex generators viz. Baseline Ramps (BRs) and Ramped Vanes (RVs). Experimental and computational investigations were conducted at Mach 2, on a cone–cylinder–flare model, with a flow turning angle of 24° at the cylinder–flare juncture. The vortex generators were placed 50 mm (10δ) upstream of the compression corner and their height was 50% of the local boundary layer thickness (h = 2.1 mm = 0.5δV G). The counter-rotating vortices emanated by Baseline Ramps (BRs), remained in close proximity to one another and interacted strongly along the upwash regions. Hence, they quickly dissipated and moved away from the model surface, causing only a nominal impact on the interaction region. Alternatively, the vortices produced by Ramped Vanes (RVs) were amply separated and therefore, largely retained their core strength and were not rapidly advected away from the near-wall region. As a result, the separation occurred a little prematurely in the upwash regions and was significantly delayed along the downwash regions. Additionally, tornado-like vortices were formed, due to the streamwise vortices interacting with the separation bubble. The delay in separation along the downwash regions, also brought the separation shock closer to the compression corner. This was accompanied by a suitable increase in the shock’s oscillating frequency, in such a way that the Strouhal number based on the interaction length (distance between the corner and the mean position of the separation shock foot) remained more or less the same.

Grid Adaptive Technique for Simulation of Scramjet Intake-Isolator at Hypersonic Speeds

Hypersonic intake is one of the major components of Scramjet engine. It compresses the incoming hypersonic flow through a series of oblique shocks as the flow passes through intake-isolator section before entering the combustion chamber, which is essential for efficient combustion. The shocks generated inside in the intake interacts with boundary layer following shock boundary layer interaction and flow separation. The separated flow blocks the flow capture area such that engine expresses unstarting phenomenon. Understanding and mitigating such flow phenomenon is a challenging task. With respect to hypersonic speeds the experimental facilities are very limited. The only alternative to solve this problem is Computational Fluid Dynamics because of its capabilities. But validation of CFD results with analytical or experimental is the foremost prerequisite to chase computational analysis. Mostly at high speeds the precision of CFD results rest on the type of grid, number of elements and turbulence model used. So, in this paper, computational analysis of hypersonic intake is carried out through designed conditions to ensure the correct CFD process is used by varying number of elements in fluid domain by grid adaptive technique using ANSYS Fluent and satisfying Y+ parameter. The domain is analysed with various turbulence models and among them SST has predicted all the flow characteristics of scramjet intake-isolator at hypersonic speeds like separation bubble, shock reattachment, cowl shock etc similar to experimental results with the help of grid adaptive technique. So, grid adaptive technique is also proposed for simulation of scramjet intake at off-design conditions.

Lifetime-based phosphor thermometry for global heat flux measurement in a hypersonic rarefied tunnel under strong background radiation

Lifetime-based phosphor thermometry using Mg4FGeO6:Mn is successfully implemented on a standard hypervelocity ballistic model 2 (HB-2 model) and an HB-2 model modified with tail fin in a hypersonic rarefied tunnel at a Mach number of 10. A novel Reynolds decomposition correction method is proposed to remove the heating device-induced strong background radiation, which is characterized by complex spatial and temporal distributions. Three methods are compared for heat flux calculation: an insitu calibration method, the Cook–Felderman method, and a double-layer numerical method. The in situ calibration method exhibits the least relative error (9.47 %), outperforming the other two methods. The global heat flux distribution at a high spatial resolution (0.25 mm/pixel) reveals the fine flow features of the models. Particularly, the modified HB-2 model exhibits a complex flow, related to the shock-boundary layer interaction, shock–shock interaction, and significant 3D effects. The presented measurement technique and data processing method can enable long-duration advanced aerothermal testing in hypersonic tunnels involving a large temperature rise.

Nozzle geometry effects on supersonic wind tunnel studies of shock–boundary-layer interactions

Many supersonic wind tunnel experiments investigate shock–boundary-layer interactions by measuring the response of tunnel wall boundary layers to an incident shock wave. To generate the supersonic flow, these facilities typically use two-dimensional contoured converging–diverging nozzles which can be arranged in two different ways. One configuration is symmetric about the centre height, whereas this symmetry plane defines the tunnel floor in the other asymmetric arrangement. In order to determine whether these nozzle configurations, which are widely thought to be equivalent, can influence experiments on shock–boundary-layer interactions, two different nozzle geometries are compared with one another in a single facility with rectangular cross section. For each setup, a full-span 8-degree wedge introduces an oblique shock to a Mach 2.5 flow. The two setups exhibit quite dissimilar behaviour, both in the corner regions and on the tunnel’s centre span, with a difference in central separation length of as much as 35% suggesting that nozzle geometry can have a profound impact on these interactions. The observed behaviour is caused by known secondary flows in the sidewall boundary layers which are driven by vertical pressure gradients in the nozzle region. The subsequent impact on the response of the floor boundary layer is consistent with expectations based on local flow momentum affecting corner separation size and on the displacement effect of this corner separation influencing the wider flow.Graphical abstract

Shock Boundary Layer Interaction and Aero-Optical Effects in a Transonic Flow over Hemisphere-on-Cylinder Turrets

Hemisphere-on-cylinder turrets are the main airborne optical platform structure. However, an unsteady shock boundary layer interaction (SBLI) would act on flow separation and turbulent wake, which causes serious aero-optical effects with high spatial and temporal frequency characteristics. In this paper, the SBLI phenomenon of a hemisphere-on-cylinder turret is recorded in a wind tunnel at Ma = 0.7 using shadowing and Mach-Zehnder interferometer measurements. Its wavefront distortion is measured using the Shack-Hartmann measurement. The detached eddy simulation (DES) based on SST k - ω turbulence model and ray-tracing methods are used to reproduce the transonic flow and optical aberration. Experiments and simulations suggest that the SBLI causes the flow to separate earlier relative to a subsonic flow over the turret. The time-averaged root-mean-square of optical path difference (OPD) over the beam aperture is 0.56 λ∼0.59 λ with λ as the wavelength, while the root-mean-square of the time-averaged OPD is about 0.45 λ. The local shock and wavefront distortion have dual peak frequencies at S t D = f D / U ∞ = 0.24 and 0.34, different from the single-peak-frequency phenomenon of a subsonic flow over turrets. Fast model decomposition of wavefront can be performed by proper orthogonal decomposition (POD) of its Zernike coefficients. The first two modes contain the shock’s reciprocating motion.

Performance evaluation of different micro vortex generators in controlling a flare-induced shock–boundary layer interaction

This study evaluates the ability of five micro vortex generator (MVG) geometries in attenuating the flow separation induced by an axisymmetric compression corner. Experimental and computational investigations were carried out at Mach 2 on a cone–cylinder–flare model, with a flow deflection angle of 24° at the cylinder/flare juncture. The MVG shapes considered were baseline ramp, trapezoidal ramp, split ramp, thick vanes and ramped vanes. A circumferential array of these MVGs having a device height (h) of 1.4 mm and an inter-device spacing of 10.5 mm (7.5 h) was introduced 50 mm upstream of the compression corner. Streamwise counter-rotating vortex pairs that originated from these devices created alternate bands of upwash and downwash regions in the incoming boundary layer, which resulted in suitable three-dimensional alterations of the separation region’s size. Furthermore, surface streamline visualizations showed that the vortices induced profound topological transformations in the separated flow structure. In the absence of MVGs, spectral analysis of the pressure signal obtained from the separation shock’s intermittent region revealed a relatively broadband dominant frequency range of 0.55 kHz–0.9 kHz. The MVGs did not cause any significant change in the dominant frequency, but made the bandwidth slightly narrower. Among the different MVG designs that were studied, the ramped vanes (RV) induced the most momentum augmentation in the near wall region and thereby caused the maximum downstream shift in the separation shock’s position.

Analytical model for predicting the length scale of shock/boundary layer interaction with curvature

In practical aerodynamic problems, curved shock/boundary layer interaction (CSBLI) is more frequently encountered than the canonical shock/boundary layer interaction (SBLI). Owing to the topological complexity of the flow field brought about by shock curvature, accurate prediction of the interaction length scale of CSBLI is a challenging task. In this work, streamwise and spanwise curvatures are introduced, in turn, with the aim of establishing an analytical model for the interaction length scale of CSBLI based on conservation of mass. The validity and universality of the model are verified, which reveal the impact of the shock curvatures on the interaction, acting as the form of non-uniformity. As an example, the CSBLI with different curvatures is compared, demonstrating that a streamwise curvature of 0.071 will bring a reduction of about 16.5% of the interaction length. The proposed method can be regarded as providing a foundation for further research on CSBLI, opening new perspectives for the investigation of SBLI flow structures.

Direct measurement of skin-friction in shock/boundary-layer interaction flows

The direct measurement of skin friction at high speed is extraordinarily complicated due to the interference from various extraneous forces and the short test-time of high-speed facilities. The present study deals with the design and performance assessment of a skin friction sensor in a hypersonic flowfield with shock-boundary layer interaction, which is a predominant high-speed flow phenomenon. The experiment was conducted in the IIT Bombay Shock Tunnel (IITB-ST); the sensor was exposed to a flow with freestream Mach number, total enthalpy, and Reynolds number of 8.4, 0.69 MJ/kg, and 2.60 × 106 m−1, respectively. The shock-boundary layer interaction was induced by an oblique shock impinged on a flat plate, with a flow turn angle of 15°. The low magnitude average skin friction could be measured with an uncertainty of ±21%. The comprehensive study conducted confirmed the suitability of the skin friction sensor for high-speed applications in impulse facilities.

Unsteady mechanisms in shock wave and boundary layer interactions over a forward-facing step

The flow over a forward-facing step (FFS) at $Ma_\infty =1.7$ and $Re_{\delta _0}=1.3718\times 10^{4}$ is investigated by well-resolved large-eddy simulation. To investigate effects of upstream flow structures and turbulence on the low-frequency dynamics of the shock wave/boundary layer interaction (SWBLI), two cases are considered: one with a laminar inflow and one with a turbulent inflow. The laminar inflow case shows signs of a rapid transition to turbulence upstream of the step, as inferred from the streamwise variation of $\langle C_f \rangle$ and the evolution of the coherent vortical structures. Nevertheless, the separation length is more than twice as large for the laminar inflow case, and the coalescence of compression waves into a separation shock is observed only for the fully turbulent inflow case. The dynamics at low and medium frequencies is characterized by a spectral analysis, where the lower frequency range is related to the unsteady separation region, and the intermediate one is associated with the shedding of shear layer vortices. For the turbulent inflow case, we furthermore use a three-dimensional dynamic mode decomposition to analyse the individual contributions of selected modes to the unsteadiness of the SWBLI. The separation shock and Görtler-like vortices, which are induced by the centrifugal forces in the separation region, are strongly correlated with the low-frequency unsteadiness in the current FFS case. Similarly as observed previously for the backward-facing steps, we observe a slightly higher non-dimensional frequency (based on the separation length) of the low-frequency mode than for SWBLI in flat plate and ramp configurations.

Evaluating the Aerodynamic Damping At Shock Wave Boundary Layer Interacting Flow Conditions with Harmonic Balance

Abstract The aerodynamic damping of a stator vane located in a rear stage of a high pressure compressor is evaluated at transonic flow conditions by numerical means. The results of a solution method based on temporal linearization around a steady RANS state are compared to results generated by relying on a state-of-the-art harmonic balance solver. It is found that the time-linearized method is not capable to reproduce the damping behavior in a sufficient manner for the majority of assessed nodal diameters. The limitations of the time-linearized method consisting in expansion around an imperfect steady RANS state, neglecting non-linear contributions and considering turbulence to be frozen at its steady state are evaluated and quantified. Furthermore, the presence of unsteady content induced by shock wave boundary layer interaction can be identified by performing a full-annulus URANS simulation based on time-integration. For a limited range of nodal diameters, this shock boundary layer interaction locks in to the assessed flutter motion and affects the damping behavior substantially. It is demonstrated that the mechanism of the shock wave boundary layer interaction can be reproduced with harmonic balance in accordance to the time-integration method. Coupling the eigenmode of interest and the identified shock boundary layer interaction via the harmonic balance method allows to predict the aerodynamic damping for the affected nodal diameters.

Unsteadiness of shock-boundary layer interactions in a Mach 2.0 supersonic turbine cascade

Shock-boundary layer interactions in a supersonic turbine cascade are investigated using a wall-resolved large eddy simulation. Results reveal that near-wall streaky structures (red color) drive the motion of the suction side separation bubble (blue color), which in turn promotes oscillations of the reattachment shock (black color) and shear layer flapping.

Atomic Layer Thermopiles: Comprehensive static calibration, comparison and application in subsonic and supersonic flows

This paper compares the calibration of different heat flux sensors in radiation- and convection-based sub- and supersonic operation. First, four heat flux sensors based on different principles: ALTP (Transverse Seebeck Effect), HFM-8E (differential-layer device), coaxial Thermocouple and a TG-2000 (circular-foil gage) are calibrated in a laser-based radiation setup. In a second step, all heat flux sensors are compared with a slug-calorimeter within a subsonic convection-dominated facility based on stagnation-point measurements of an impinging hot air jet. The obtained results indicate that a sensitivity transfer between a radiative calibrated sensor used in a mainly convective environment is not always possible and can lead to significant, systematic errors. Calibration in a subsonic shear flow with thermocouple readings demonstrate small divergence from the manufacturer provided sensitivity. Finally, supersonic testing with high frequency shock-boundary layer interactions highlight the need for ALTP rather than the conventional use of thermocouple arrays to resolve the high frequency phenomena associated with shock boundary layer interactions.

Study on the mechanism of shock wave and boundary layer interaction control using high-frequency pulsed arc discharge plasma

This paper studies the response characteristics of shock wave and boundary layer interaction (SWBLI) controlled by high-frequency pulsed arc discharge (PAD) in a Mach 2.5 flow. The dynamic evolution of SWBLI disturbed by arc plasma energy deposition was captured, and the controlling mechanism under different exciting power and frequency was explored. The results showed that the blast wave induced by PADs had a strong impact on SWBLI structures and distorted the separation shock wave. During the downstream propagation, the controlling gas bubbles (CGBs) delivered a continuous thermal excitation to the boundary layer and reached the maximum penetration depth near the semi-cylinder. The arc discharge in the SWBLI region induced larger energy deposition, which made the heating zone obtain the highest initial temperature and longest heating duration. Under the plasma condition of 1 × 1011 W/m3/15 kHz, both the upstream part of the shear layer and the foot portion of the reattachment shock wave were removed. When setting the excitation to 2.5 × 1010 W/m3/60 kHz, a thermal exciting surface of merged CGBs was formed and the separation shock wave was completely replaced by an equivalent compression-wave system. A better drag-reduction effect on the flow field would be produced by the actuator with an increased operating power or frequency, and a drag reduction rate of nearly 25.5% was achieved under the 2.5 × 1010 W/m3/60 kHz control condition.

Controlling transonic shock–boundary layer interactions over a natural laminar flow airfoil by vortical and thermal excitation

Investigations have been performed via implicit large eddy simulations to study the overall effects of exciting a flow field by thermal (wall-heating and wall-cooling) and vortical (with high and low frequencies) actuation. The actuator is placed on the suction surface of a natural laminar flow (SHM-1) airfoil having an angle of attack of α=0.38° (cruise setting). Oncoming flow has a Mach number of 0.72, and a Reynolds number based on a chord of Re=16.2×106, for which a complex shock system is formed on the suction surface. Vorticity dynamics of the flow is studied using time series of vorticity at different locations above the suction surface and instantaneous contour plots of vorticity in the domain. An inspection of the flow using snapshots of ∇ρ and ∇(ρT) is done to characterize the numerical schlieren. The comparative effects of the various forms of excitation on the shock–boundary layer interactions (SBLI) have been analyzed using time series of the magnitudes of ∇ρ across the identified shock structures from numerical schlieren snapshots. Also, the role of the frequency of imposed vortical actuation has been studied using vorticity and Mach contours for a comparative understanding of the control of the SBLI.

Swept shock wave/boundary layer interaction control based on surface arc plasma

Swept shock wave/boundary layer interactions occur widely in the internal and external flows of supersonic and hypersonic aircraft, which can seriously affect aircraft stability and engine intake efficiency. Based on the conventional S–A turbulence model, this study investigates surface arc plasma actuation for regulating swept shock wave/boundary layer interactions at Mach 2.95 to explore the ability and the three-dimensional shock wave/boundary layer interactions control method of plasma actuation. First, the flow control effect is explored in terms of indirect control by applying actuation in the upstream boundary layer or in front of the separation line, and in terms of direct control by applying actuation in the separation region. These three methods all achieve clear control effects. Control results show that the first method is more effective in regulating the wall pressure and friction coefficient and can improve the friction and heat transfer of the wall in a wide range of flow direction and cone direction. The second method is more effective in regulating separated shock waves. The third aspect is more effective in regulating the reattachment region. The associated control mechanisms are then refined. The control effects of the first control method depend on the transmission of vortices, those of the second are based on the virtual surface generated by actuation, and those of the third rely on energy injection. Finally, the application scenarios of the different control methods are determined according to the flow control requirements of aircraft and the corresponding control mechanisms. This study provides a reference method for solving more complex three-dimensional shock boundary layer interaction problems.

Boundary-Layer Instabilities over a Cone–Cylinder–Flare Model at Mach 6

Computations are performed to investigate the convective and global boundary-layer instabilities over a sharp cone–cylinder–flare model at zero-degree angle of attack. The model geometry and the flow conditions are selected to match the experiments conducted in the Boeing/AFOSR Mach 6 Quiet Tunnel at Purdue University. The geometry consists of a nominally sharp 5 deg half-angle cone, followed by a cylindrical segment, and then a 10 deg flare. Additionally, flare half angles equal to 8 and 12 deg and nosetip radii equal to 0.1, 1, and 5 mm are studied. An axisymmetric separation bubble is generated as a result of the laminar shock–boundary-layer interaction in the cylinder–flare region. The comparison of the laminar flow solution and the schlieren images shows a remarkable agreement between the respective locations of both the boundary-layer edge and the reattachment shock. The linear amplification of first and second Mack mode instabilities that begin to amplify in the cone region are computed with a combination of the parabolized stability equations and the harmonic linearized Navier–Stokes equations. The predicted frequency spectra of the surface pressure fluctuations associated with both planar and oblique instability waves capture the distinct lobes within the disturbance amplification spectra measured with PCB® and Kulite sensors. To our knowledge, this represents the first successful comparison between convective instability analysis and measured surface pressure fluctuations for a hypersonic configuration with a separation bubble. Finally, the global instability analysis shows that the laminar flow becomes supercritical for flare half angles larger than 8 deg.

Dynamic interaction between shock wave turbulent boundary layer and flexible panel

The high-speed fluid–structure interaction (FSI) occurring between a skin panel and a turbulent shock-boundary layer interaction (SBLI) are investigated with large eddy simulations (LES). The SBLI is comprised of a Mach 4 and unit Reynolds number of 2.375×107 (/m) flow interacting with an oblique shock of strength p3/p1≈8.5 and angle σ≈30deg. The inviscid point of impingement on the flexible panel occurs near the mid-chord length. The fully clamped rectangular panel has the aspect ratio and non-dimensional panel thickness of b/a=1.377 and h/a=0.003, respectively. The strong SBLI induces a large flow separation of length Lsep≈0.7a (or equivalently ≈23δin) with a wide range of turbulent scales, including the characteristic low-frequency unsteadiness. Fluid and structural responses are compared across different classes of interaction, namely: rigid surface; one-way application of SBLI loads to the structure; and two-way interaction between the flow and structure. The interplay between the SBLI and flexible panel is elucidated using detailed analysis that includes proper orthogonal decomposition (POD), dynamic mode decomposition (DMD), and the recently developed Lagrangian modal analysis (LMA) for the deforming meshes. The large panel deflections (∼1h) result in structural non-linearity associated with the coupling between bending and stretching of the panel, manifested in terms of increased modal frequencies of the panel deflection.

On the scaling of three-dimensional shock-induced separated flow due to protuberances

Supersonic flow over three-dimensional bodies protruding out of the turbulent boundary layer was investigated by means of experiments and numerical computations. A parametric study was performed by varying the shape and dimensions of the protuberance, as well as the freestream Mach numbers (1.5, 2, 2.5, 2.89, and 3.5). Surface streak line visualization, surface pressure measurements, and time-resolved Schlieren visualization were employed along with Reynolds-averaged Navier–Stokes computations to elicit the complex flow features such as the separation line, shock pattern, and the horseshoe vortex, which greatly influence the flow dynamics in the separated region. The rise in surface pressure at mid-span due to separation (plateau pressure) was dependent only on the incoming flow parameters and independent of protuberance geometry. The two-dimensional free interaction theory, applied for normal shock-induced separation, closely predicts the mid-span plateau pressure. Although protuberances are of varying shapes and dimensions, the inviscid bow shock (obtained from Euler computations) provided generalized scales, whose effects on shock boundary layer interactions are analyzed. The radius of curvature of the inviscid shock on the wall plane at the nose, which is theoretically related to the local second derivative (along the shock) of pressure jump, was found to be a determining parameter of mid-span separation length (Lsep). Since the spanwise distance of the sonic point on the inviscid shock was found to be strongly correlated with its nose radius of curvature, it follows that the “strong” portion of the inviscid bow shock fixes the mid-span separation location. These observations concerning mid-span plateau pressure, and the role of strong shock portion in fixing mid-span separation, suggest that the Lsep shall be predicted from a modification of the scaling laws for the length of plateau pressure region in two-dimensional shock boundary layer interaction, with the inclusion of a spanwise relieving effect. A correlation is obtained relating the Lsep with various incoming flow parameters and inviscid shock nose radius. The mid-span vortex core position was found to be linearly related to the Lsep. The radius of curvature of the separation shock is, however, found to be influenced by the entire inviscid shock, including the “weak” portion.

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.

Control of flow separation over an axisymmetric flared body using ramped vanes

Experiments were performed at Mach 2 to assess the effectiveness of Ramped Vanes (RVs) in delaying the onset of shock-induced flow separation due to an axisymmetric compression corner. A cone–cylinder-flare model with a 24° flow deflection angle at the cylinder/flare juncture was selected for this study. A single row of RVs was placed 50 mm (10δ) upstream of the compression corner, to manipulate the incoming boundary layer and thereby bring alterations to the shock–boundary layer interaction flowfield. RVs with three sub-boundary layer heights viz. 1.4 mm, 2.1 mm and 2.8 mm (h/δRV=0.33, 0.50 and 0.67 respectively) were employed and their separation control performance was compared. Counter-rotating vortices originated from these devices, generating alternate upwash and downwash bands within the boundary layer. When the shorter devices were used, early and delayed onset of separation were noticed along the upwash and downwash regions respectively. However, as the devices grew taller, the increase in separation length along the upwash region seemed to diminish. Ultimately, when the tallest devices (h=2.8 mm) were introduced, the entire corrugated separation line was successfully pushed downstream of the uncontrolled separation line. Besides instigating these variations in the separation scale, the vortices also caused topological changes within the separation region. But, despite achieving a substantial delay in separation, the RVs were unable to shift the separation shock’s oscillations to higher frequencies.