Incoming Boundary Layer Thickness Research Articles

Effect of incoming flow conditions on air lubrication regimes

Different air phase regimes are formed by controlled air injection in a spatially developing flat plate turbulent boundary layer (TBL). The air is introduced via a slot type injector without the use of a backward-facing step or cavitator upstream of the air injection position. The effect of different incoming liquid flow characteristics on the different regimes is investigated by varying both the liquid freestream velocity and the incoming TBL thickness. The latter is realized through changing the position of the air injection along the length of the water tunnel facility. That resulted in a downstream distance based Reynolds number from 1 to 5 million. Three different air phase regimes are identified under different air flow rates and freestream velocities: the bubbly regime, the transitional, and the air layer regime. The morphological differences of each one are described and quantitative analysis is performed based on the non-wetted area in each condition. The incoming TBL as well as the flow around the air layer are measured with planar particle image velocimetry. The latter enabled the determination of the air layer thickness. In addition, the ratio of the air layer to the incoming boundary layer thickness tair/δ is also calculated (≈ 0.04 – 0.5). This is a significant dimensionless parameter for scaling, which indicates the extent to which the air layer is embedded within the incoming TBL. Depending on the incoming flow conditions, a two or three branch air layer is formed. The length of the air layer is found to increase with increasing liquid freestream velocities. A good agreement between the air layer length and a half gravity wave predicted by the dispersion relation is found. An increase of the air layer length is observed with a decreasing incoming TBL thickness. This is attributed to a decrease in the local mean velocity at the air–water interface due to the TBL growth. Finally, increasing the incoming TBL thickness delays the onset of the air layer regime.

Study on the Control Mechanism of DBD-PA on Near Wall Flow

Based on the Shyy plasma simulation model, this paper studied the control mechanism of positive and reverse DBD-PA (dielectric barrier discharge plasma actuator) on near wall flow, and discussed the influence of incoming boundary layer thickness and parallel double DBD-PA on control effect. The results showed that: the positive DBD-PA had an obvious acceleration effect on the near wall flow field, and the acceleration effect first increased and then decreased; the reverse DBD-PA had an obvious deceleration effect on the near wall flow field, and the deceleration effect decreased gradually; increasing the thickness of the incoming boundary layer can improve the control effect, but the improvement effect is gradually reduced; when parallel double DBD-PA, the control effect of positive and reverse DBD-PA is improved by 88.7 % and 62.5 % respectively.

Coexistence of different mechanisms underlying the dynamics of supersonic turbulent flow over a compression ramp

Supersonic turbulent flow over a compression ramp is studied using wall-resolved large eddy simulation with a freestream Mach number of 2.95 and a Reynolds number [based on δ0: the thickness of incoming turbulent boundary layer (TBL)] of 63 560. The unsteady dynamics of the present shock wave/turbulent boundary layer interaction (STBLI) flow are investigated by using dynamic mode decomposition techniques, linear and nonlinear disambiguation optimization, local stability analysis (LSA), and global stability analysis (GSA). By analyzing the dynamic system for the STBLI flow, three dynamically important modes with characteristic spanwise wavelengths of 2δ0, 3δ0, and 6δ0 are captured. The 2δ0 mode approximates the spanwise scale of the Görtler-like vortices and Görtler mode of LSA, suggesting the presence of Görtler instability, which is believed to be related to the unsteady motion of streaks downstream of reattachment in the flow. The features of the 3δ0 mode are also observed in large-scale motions of the incoming TBL, implying the existence of a convective mechanism that is excited and maintained by such motions. Additionally, the GSA results show the most unstable mode features a spanwise wavelength of around 6δ0, indicating the existence of global instability that is believed to be related to the oscillating motion of separation shock. The coexistence of these three mechanisms is confirmed. Discussions on the above findings provide an interpretation for low-frequency unsteadiness that the unsteadiness of surface streaks results from the combined effects of the Görtler instability near flow reattachment and the convection of large-scale motions in the incoming boundary layer, while the low-frequency shock motion may be related to a global mode driven by upstream disturbances.

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.

Aeroacoustics of turbulent flow over a forward–backward facing step

This paper presents a combined experimental and numerical study on the aeroacoustics of low Mach number turbulent flow over a forward–backward facing step (FBS). The height of the step is fixed at 50% of the incoming boundary layer thickness. Multiple lengths of the step are considered with aspect ratios (step length to height) equal to 1,2,2.4,4, and 8. The Reynolds number based on the step height ranges from 1.2×104 to 2.8×104 in both experiments and simulations. Wall pressure measurement results show that the largest wall pressure fluctuations occur slightly downstream of the step leading edge, and the downstream flow field is influenced by the leading edge flow disturbance. Large-Eddy Simulation (LES) is performed to visualize the instantaneous and time-averaged flow field. Large-scale vortices formed at the leading edge are found to shed downstream directly when the step length is shorter than the shear layer reattachment length, which leads to lower level wall pressure fluctuations on the top surface. Spectral proper orthogonal decomposition (SPOD) is applied to identify the dynamic behaviours of coherent structures downstream of the step leading edge. Using simulation data, the fluctuating wall pressure spectra are also calculated via Poisson’s equation. The results indicate that both turbulence-mean-shear interaction and turbulence–turbulence interaction contribute to the wall pressure fluctuations in the on-step reattachment region. Aeroacoustic beamforming results show that the dominant noise source is located at the step leading edge, where the highest level of fluctuating wall pressure is recorded. Meanwhile, the integrated far-field acoustic spectra show that the levels of FBS noise are lower when there is no flow reattachment occurring on the top surface. A new scaling law for FBS sound spectra is proposed to address the influence of the step aspect ratio on noise production.

Response of hypersonic compression corner flow to upstream disturbances

Hypersonic flow over a two-dimensional compression corner is investigated using computational fluid dynamics and global resolvent analysis in this study, which is a continuation of the work done by Hao et al. (J. Fluid Mech., vol. 919, 2021, p. A4). The same baseline free-stream conditions with a Mach number of 7.7 and a unit Reynolds number of 4.2 × 106 m−1 are considered. The ramp angles range from 0° (equivalent to a flat plate) to 12° (slightly below the critical angle of global instability). During this process, the base flow evolves from no separation to incipient separation to large separation. The resolvent analysis reveals that the optimal response to upstream disturbances localised near the leading edge is in the form of steady streamwise streaks for all interaction strengths, which arise from transient growth in the flat-plate boundary layer due to the lift-up mechanism and significantly grow near the corner due to the Görtler instability. The most amplified spanwise wavelength decreases as the ramp angle is increased and scales with the incoming boundary-layer thickness.

Investigations of shock–boundary layer interaction dynamics using high-bandwidth pressure field imaging

The large-scale pulsations of shock-induced separation with length scale that significantly exceeds the incoming boundary layer thickness are investigated. The shock–boundary layer interaction (SBLI) unit is generated by an inward-turning axisymmetric compression ramp at an inflow Mach number of 2.5. A substantial region surrounding the centre azimuth exhibited mean and dynamic flow features that are consistent with two-dimensional separation. Two-dimensional highly resolved maps of surface pressure field are obtained using fast-response pressure-sensitive paint fluorescence imaging at 40 kHz repetition rate. The measurement domain covered significant regions of the incoming boundary layer through the relaxing boundary layer downstream of the reattachment as well as over 25 boundary layer thicknesses in the azimuthal direction. These measurements provide new insights into the spanwise coupling of the SBLI unit in addition to its inherent dynamics. The power spectral density (PSD) of the centreline pressure exhibits very good agreement with theoretical models and complementary measurements using fast-response pressure transducers, which served to validate the pressure field measurements. Detailed examination of the PSD reveals strong agreement with the literature, which includes the peak Strouhal number of the separation and reattachment shock motions as well as the downward frequency shift along the separation bubble. Furthermore, the pressure fluctuation maps reveal streamwise-elongated structures just downstream of the ramp leading edge that persist well downstream of the reattachment. A time sequence of conditional average pressure fluctuation maps is constructed surrounding isolated pressure excursions in the intermittent region. This sequence, along with two-point cross correlation analysis, provides critical information about the flow processes that drive the separation bubble pulsations in the SBLI units with large separation scales. Overall, the imbalance in the mass within the separation bubble appears to be the critical mechanism that drives the separation bubble pulsations. Furthermore, the pressure perturbations originating at azimuthally offset locations are also observed to influence the separation bubble dynamics.

Analysis of the Effect of the Leading-Edge Vortex Structure on Unsteady Secondary Flow at the Endwall of a High-Lift Low-Pressure Turbine

The horseshoe vortex system at the leading edge of a low-pressure turbine (LPT) has unsteady flow characteristics, and the flow field in the downstream cascade channel also has unsteady characteristics. In this study, CFD simulations are performed with the help of commercial software CFX, and experimental checks are performed using a fan-shaped cascade test bench to investigate the flow characteristics of the endwall of the PACKB blade type of a high-lift LPT under the same incoming Reynolds number and two incoming boundary layer thickness conditions. The obtained results show that there are different flow alteration characteristics and alteration frequencies of the horseshoe vortex system, among which the vortex system with a thicker boundary layer is larger in size and less spaced from each other, which is more likely to induce the fusion of vortex systems. The centrifugal instability causes the instability of the horseshoe vortex system, and the instability frequency is inversely proportional to the thickness of the boundary layer. With two inlet boundary layers, the instability frequencies of the vortex system are 125 Hz and 175 Hz, respectively, and the ratio of the frequency to the thickness of the boundary layer is reciprocal to each other. The stimulation effect of the unstable horseshoe vortex system on the downstream secondary flow intensity is greater than that of the steady state. The thin boundary layer case generates a greater unsteady loss in the cascade channel than the thick boundary layer case due to the poor stability of the vortex system.

Gap ratio effects on the coherent structures surrounding a near-wall square cylinder

Turbulent flow around a square cylinder offset at different heights from an upstream rough wall are investigated using time-resolved particle image velocimetry. The Reynolds number based on the freestream velocity and cylinder height was 12750 and the incoming turbulent boundary layer thickness was 7.2 times the cylinder height. The examined gaps beneath the cylinder were 0.0, 0.3, 0.5, 1.0, 2.0, 4.0, and 8.0 times the cylinder height. The results show that the mean recirculation length remains relatively constant for cases with a gap size greater than or equal to 1.0 times the cylinder height. Flow acceleration through the gap under the cylinder is strongest with a gap size of 2.0 times cylinder height. Downstream of the cylinder, the vertical velocity fluctuations contribute strongly to the production of turbulent kinetic energy by interacting with mean wall-normal velocity gradient in the vertical direction. The dominant frequency embedded in the incoming turbulent boundary layer persists in the wake of the cylinder when the gap size is less than 1.0 cylinder height. Kelvin-Helmholtz vortices originating from the top leading edge of cylinder occur at Strouhal numbers of 2.77-3.11 for the cases with gap ratio ranging from 0.3-8.0 cylinder heights, while spectral migration towards lower frequencies is observed for cases with gap sizes equal to 0.3 and 0.5 cylinder heights. The spatial trajectory of top Kelvin-Helmholtz vortices extends further away from the cylinder as the gap size increases. Spectral proper orthogonal decomposition analysis shows that the von Kármán vortex shedding structures possess the highest spectral energy for all offset cases despite regular vortex shedding being disrupted when the gap size is 0.3 times the cylinder height.

Downstream influence of turbulent flow past vortex generators

Abstract Vortex generators (VGs) are typically positioned upstream of a boundary layer separation region. Their effectiveness depends on incoming flow conditions (Mach number, Reynolds number, and boundary layer characteristics), geometrical configuration (vane or ramp, height, width, and angle of incidence), and spacing. Device-induced vortices and following decay allow VGs to be used as a passive control device. This study uses pressure-sensitive paint to determine the global surface pressure pattern for a flat plate flow in the presence of VGs (counter-rotating vanes, co-rotating vanes and ramps). The freestream Mach number is 0.64 and 0.83. The ratio between the height of the VGs and the incoming boundary layer thickness is 0.2, 0.5, and 1.0. The standard deviation in the spanwise pressure in the streamwise direction is used to determine the downstream influence of turbulent flow past VGs. Increasing the height of VGs causes device-induced vortices to propagate farther downstream.

Effect of Swept Shockwave Boundary-Layer Interaction Strength on Surface Skin Friction

This investigation examines the flowfield topology beneath the single fin-generated shockwave boundary-layer interaction as a function of shock interaction strength and incoming boundary-layer thickness. A systematic application of oil film interferometry produced an experimental reconstruction of global skin friction maps for various Mach number and fin angle combinations. To determine the effect of boundary-layer thickness, data from the current study were compared against existing experimental data sets with thinner incoming boundary layers. Results for identically matching flow/geometric conditions show a good agreement between trends seen across the conical region of the interaction and exhibit skin friction coefficient peaks in the same position. Cases corresponding to a relatively thicker boundary layer showed higher skin friction coefficient values along the interaction periphery and upstream regions. An angular scaling was applied to align the inviscid shocks of data sets that matched interaction strength, and the results show similar behavior. The peak skin friction coefficient shows a monotonic increase with increasing shock strength. Conversely, the upstream normalized peak skin friction coefficient depends on boundary-layer thickness and shock strength. These findings indicate that boundary-layer effects are limited to the interaction periphery, whereas viscous effects driven by shockwave–boundary-layer interactions dominate the interior region.

Influence of isolator section on the shock augmented mixing in SCRAMJET engine

The paper aims to numerically study the effect of isolator design on shock-augmented mixing in a generic supersonic Ramjet (SCRAMJET) engine and reduce what would be a very complicated thermodynamics+aerodynamic optimization in the design to a simpler aerodynamic one. Changing the length of the isolator section effectively varies the thickness of the boundary layer that is formed in the isolator, which then enters into the SCRAMJET combustor. The study is conducted on a three-dimensional geometry using the Menter's SST model for turbulence on an in-house unstructured grid RANS solver. In this paper, the authors quantify the effect of the incoming boundary layer thickness on the shock-augment air-fuel mixing in the SCRAMJET combustor. It is observed that for the maximum variation in the isolator length, the normalized mixing volume only improves by 7%, which is only a marginally better mixing of air and fuel obtained as with an increase in the isolator length. The authors conclude that the Prandtl-Meyer expansion fan (PMEF) formed at the flame-holder effectively isolates the combustor from the isolator section to a degree that the isolator dimensions can therefore be chosen entirely on considerations of flow homogenization of inlet air before it enters the combustor section of the SCRAMJET engine. Although, the paper analyses a generic SCRAMJET with specific dimensions, the results are more generally applicable than for the specific design actually considered in the paper. Thus this study is also relevant to SCRAMJET engines that may differ radically from our considered design in form and dimensions.

Quantitative flow diagnostics of shock trains by rainbow schlieren deflectometry

AbstractA new measurement technique to reconstruct the density field of the shock-wave/boundary-layer interaction (SWBLI) in a confined duct is proposed. With this technique, it is possible to quantitatively capture in detail the structures of the density field both in the regions of the shock-systems in the central core and boundary-layer flows near the duct wall concurrently. The novel feature of the proposed technique is to make use of the schlieren images with the rainbow filters of the vertical and horizontal cutoff settings and then to reconstruct the two-dimensional density field integrated over the line-of-sight direction using the corresponding filter calibration curves. The proposed technique is applied for the first time to a shock train in a constant-area straight duct under the upstream condition of the shock train: the freestream Mach number is 1.42, the incoming boundary layer thickness normalised by the duct half height is 0.175, and the corresponding unit Reynolds number $Re/m$ is $2.99 \times 10^7$ m-1. The calculated isopycnic field depicts the streamwise and transverse density variations inside the shock train, the mixing region after the shock train, and the boundary-layer of the interaction region. This technique is shown to be capable of identifying the locations of shocks in a shock train more precisely than a conventional approach measuring the static pressure distribution along the duct wall. In addition, various quantitative visual representations such as a shadowgraphy and a bright-field schlieren can be extracted from the density field acquired by the present approach, and the spatial evolution of the shape and strength of each shock constituting the shock train as well as the boundary layer flow properties can be quantitatively clarified.

Investigation on conical separation vortex generated by swept shock wave/turbulent boundary layer interaction

To study the effects of Mach number, Reynolds number and deflection angle on the conical separation vortex generated by swept shock wave/turbulent boundary layer interaction , the experimental and numerical simulation methods are used. The Mach numbers are Ma=2.95 and Ma=4. The deflection angles are 12°, 15° and 18°. The incoming boundary layer thickness δ0.99 is 4.6 mm, 23 mm and 115 mm. The evolution characteristics of the vortex, including the vortex core position, the vortex area and the vortex intensity, and the effects of Mach number, Reynolds number and deflection angle on them are analyzed. The results show that the core trajectory of the vortex is distributed along the rear shock foot line of the λ shock structure. The Reynolds number is the main factor affecting the separation vortex core. The vortex area increases linearly along the direction of the inviscid shock, and the Reynolds number and the Mach number both are the factors affecting the growth rate of the separation vortex. The vortex intensity increases linearly along the direction of the inviscid shock. The vortex intensity is mainly affected by the Mach number and the deflection angle, but is not affected by the Reynolds number.

Flow structure of pitched and skewed microjet vortex generator array

To understand the micro-vortex structure of Microjet Vortex Generators, a plate with pitched and skewed jet holes was established. Simulations using SST-DDES method based on WENO scheme were conducted to analyse the effect of jet pressure, hole interval and incoming boundary layer thickness on the flow structure. The pitch angle and skew angle are 55 degrees and 45 degrees respectively. The total pressure of microjets ranges from 2.5 to 3.5 times the mainstream pressure. The interval of microjets ranges from 6.75 to 13.5 diameters of the hole, and the thickness of incoming boundary layer ranges from 2 to 4 diameters of the hole. The flow field is dominated by three vortices, including a main vortex and two secondary vortices which are significantly influenced by jet pressure and interval. The interaction between secondary vortices and main vortex plays an important role in the process of boundary layer mixing.

Fluid–structural coupling of an impinging shock–turbulent boundary layer interaction at Mach 3 over a flexible panel

We present high-fidelity numerical simulations of the interaction of an oblique shock impinging on the turbulent boundary layer developed over a rectangular flexible panel, replicating wind tunnel experiments by Daub et al. (AIAA Journal, vol. 54, 2016, pp. 670–678). The incoming free-stream Mach and unit Reynolds numbers are $M_{\infty } = 3$ and $Re_{\infty }=49.4\times 10^6 {\rm m}^{-1}$ , respectively. The reference boundary layer thickness upstream of the interaction with the shock is $\delta _0 = 4$ mm. The oblique shock is generated with a rotating wedge initially parallel to the flow that increases the deflection angle up to $\theta _{{max}} = 17.5^{\circ }$ within approximately $15$ ms. A loosely coupled partitioned flow–structure interaction simulation methodology is used, combining a finite-volume flow solver of the compressible wall-modelled large-eddy simulation equations, an isoparametric finite-element solid mechanics solver and a spring-system-based mesh deformation solver. Simulations are conducted with rigid and flexible panels, and the results compared to elucidate the effects of panel flexibility on the interaction. Three-dimensional effects are evaluated by conducting simulations with both full ( $50 \delta _0$ ) and reduced ( $5\delta _0$ ) spanwise panel width, the latter enforcing spanwise periodicity. Panel flexibility is found to increase the separation bubble size and modify its spectral dynamics. Time- and spanwise-averaged streamwise profiles of the wall pressure exhibit a drop over the flexible panel prior to the interaction and a reduced peak pressure in comparison with the rigid case. Spectral analyses of wall pressure data indicate that the low-frequency motions have a similar spectral distribution for the rigid and flexible cases, but the flexible case shows a wider region dominated by low-frequency motions and traces of the panel vibration on the wall pressure signal. The sensitivity of the interaction to small variations in the wedge extent and incoming boundary layer thickness is evaluated. Predictions obtained from lower-fidelity modelling simplifications are also assessed.

Effect of incoming boundary layer thickness on the flow dynamics of a square finite wall-mounted cylinder

This paper investigates the flow dynamics around a square finite wall-mounted square cylinder of two different aspect ratios (AR=H/D=6 and 10, H and D are the height and width, respectively) under six boundary layer thickness (δ/D=0.5–4.0, δ is the boundary layer thickness) using the improved delayed detached eddy simulation turbulence model. The Reynolds number based on the square cylinder width and incoming velocity is Re=1.4×104. The aerodynamic force, surface mean and fluctuating pressure, wake closure length, Reynolds stresses, turbulence production, and the anisotropy of turbulence are analyzed in detail to investigate the effect of AR and δ/D combinations. The Q-criterion vortex identification method is adopted to identify and classify the vortex structures of different AR and δ/D combinations. The parametric diagram based on AR and δ/D to predict the dipole/quadrupole wake structure proposed in Yauwenas et al. [“The effect of aspect ratio on the wake structure of finite wall-mounted square cylinders,” J. Fluid Mech. 875, 929–960 (2019)] is refined based on results of the present study (quadrupole wake for AR=6 with δ/D≥1.5 and AR=10 with δ/D≥1.0). The instantaneous space–time pressure distribution along the whole span proves the existence of cellular shedding (AR=10 for δ/D=0.5–2.0), and the spectral proper orthogonal decomposition technique is adopted to investigate the cellular shedding phenomenon and its distribution. With the increase in δ/D from 2.0 to 4.0, the cellular shedding process is absent, implying the inhibitory effect of increasing boundary layer thickness on the cellular shedding.

Linear and Nonlinear Flow Analysis of Elements of a Supersonic Inlet

Boundary layer development in a supersonic inlet is studied, focusing on the flow over the compression ramp and the ramp–isolator junction. Two designs of the junction, a backward-facing step (BFS) and a backward-facing notch (BFN), are first evaluated to identify three linear frequency bands (high, medium, and low) of the shear layer instabilities. Compared with BFS, the laminar flow and linear instabilities of BFN are largely insensitive to the incoming boundary layer thickness and show relatively weaker low-frequency unsteadiness. Subsequent linear analysis of the combined ramp and BFN junction identifies a two-dimensional Kelvin–Helmholtz (K-H) instability within the ramp separation, the spectral signature of which extends further downstream. Corresponding large-eddy simulation (LES) predicts transition over the ramp and a turbulized boundary layer at the BFN junction. Modal analysis of LES shows that transition is primarily driven by the breakdown of Görtler vortices due to secondary instabilities, which are tailored by the two-dimensional K-H mode of the ramp separation. The linear analysis provides a reasonable estimate of the pre-transition LES spectrum. The topology of the BFN junction includes streamwise vortices within the turbulent separation bubble, and low-frequency convecting structures further downstream. Statistical parameters from LES indicate strong non-equilibrium effects in the boundary layer post-transition.

Effect of Incoming Boundary-Layer Characteristics on Performance of a Distributed Propulsion System

An experimental study is conducted to advance the understanding of flow physics associated with a boundary-layer ingesting, distributed propulsion system. The influence of incoming boundary-layer thickness on the performance of the system is examined. The propulsion model, integrated with electrical fans, is mounted on a flat plate and tested at subsonic speeds. Detailed characterization of the incoming boundary layer and the downstream flowfield is performed using hot-wire anemometry and static pressure measurements. Modification of the boundary layer is achieved by placing trip rods of two different diameters near the leading edge of the flat plate. Overall performance of the system is analyzed by estimating thrust, flow power, and input power to the fans. Ingestion of a thicker boundary layer is found to result in an increase in thrust and net flow power.

Effects of Reynolds Number on Swept Shock-Wave/Boundary-Layer Interactions

The current experimental investigation explores the viscous/inviscid interaction of a supersonic turbulent boundary layer and a sharp fin-generated shock wave. The study examines the effect of Reynolds numbers on the fin-generated shock-wave/boundary-layer interactions (SBLIs) and the practical limits imposed on such scaling due to different test facilities. The present study involves a parametric variation of the incoming Mach number, fin angle of attack, and a wide range of unit Reynolds numbers in combination with advanced flowfield diagnostics. The results indicate that the interaction footprint is distinctly affected by the Reynolds number, though the effects are more prominent near the fin/surface junction and the outer edges of the interaction region. The results suggest that the effect of the Reynolds number on the SBLI is significantly reduced at higher Reynolds numbers. This sensitivity to Reynolds number is a function of the fin angle of incidence and freestream Mach number, with weaker strength interactions being more receptive to Reynolds number variation. Identical measurements carried out in facilities of different size compare favorably throughout the interaction region with Reynolds-based scaling; however, incoming boundary-layer thickness imposes limitations on the extent of the inception region and the onset of finite fin effects.