Compression Ramp Research Articles

Amplification of turbulent kinetic energy and temperature fluctuation in a hypersonic turbulent boundary layer over a compression ramp

In this paper, direct numerical simulations in a Mach 6.0 hypersonic turbulent boundary layer over a 30 ° compression ramp are performed. The influence of shock wave/boundary layer interactions on the amplification of turbulent kinetic energy (TKE) and temperature fluctuation (TF) is explored, to provide an insight into the physical mechanism. In the initial part of the interaction region before the detachment of the shear layer, the amplification of the TKE and TF is found, via a frequency spectrum analysis, to be closely related to the low-frequency unsteadiness of the shock wave. Once the free shear layer is established, the shear component of the TKE production defined in the shear layer coordinate appears to act as the main contributor for the TKE amplification, owing to the mixing layer turbulence and the resultant Kelvin–Helmholtz instability. This is consistent with the result from the spectrum analysis that the TKE and TF amplification and their streamwise evolution are dominated by the spectral energy in the median-frequency range, arising from the mixing layer turbulence. As the flow moves downstream along the shock wave, the high-frequency spectral energy content of TF shows a decreasing trend, while the low-frequency spectral energy tends to increase gradually, implying that the shock wave low-frequency unsteadiness exists not only in the initial stage of the interaction region but also around the main shock wave. Under the combined influence of the shock wave intensity and interaction intensity, the median-frequency content appears to weaken first and then tends to increase before decreasing again. The variation amplitude appears to be small and generally dominates the distribution of the TF intensity.

Large eddy simulation of shock wave/turbulent boundary layer interaction under incipient and fully separated conditions

Large eddy simulations of shock wave/turbulent boundary layer interaction on a compression ramp at the Mach number M a ∞ = 5 and Reynolds number R e ∞ = 14 000 are performed to investigate the impact of the incipient and fully separated conditions on the development of the flow field. The quasi-dynamic subgrid-scale kinetic energy equation model, which combines the benefits of the gradient model with the eddy-viscosity model, has been applied. Compared with the previous experimental and numerical results, the simulation was validated. The flow structures, turbulence properties, vortex structures, and low-frequency unsteadiness are all investigated. The flow field of the incipient separation is attached and rarely impacted by shock. An evident separation bubble and localized high wall temperatures in fully separated flow are caused by the separation shock's significant reverse pressure gradient. The Reynolds stress components exhibit significant amplification in both cases, and the peak outward shifts from the near-wall region to the center of the free shear layer. Turbulent kinetic energy terms were analyzed, and the two scenarios show a significant difference. The power spectral density of the wall pressure fluctuations shows that the low-frequency motion of the incipient separation is not apparent relative to the fully separated flow.

Stability of hypersonic flow over a curved compression ramp

In this work, the stability of hypersonic flow over a curved compression ramp is studied using several stability analysis tools and direct numerical simulations (DNS). The free-stream Mach number and the unit Reynolds number are 7.7 and $4.2 \times 10^6$ m $^{-1}$ , respectively. Corner rounding is considered to alter the separation bubble flow so as to suppress the intrinsic instability of the compression-ramp flow. The variation of intrinsic instability is confirmed by global stability analysis. Subsequently, resolvent analysis is employed to examine the response of intrinsically stable flows to external disturbances. It is shown that the considered flows strongly amplify low-frequency streamwise streaks with a preferential spanwise wavelength. This result is verified using DNS by introducing a random forcing upstream of the separation point. Furthermore, both resolvent analysis and DNS demonstrate that the separation bubble contributes little to the selection of the spanwise wavelength of streamwise streaks. The combined effects of convective and intrinsic instabilities are also explored using DNS. A better agreement with experimental data is achieved after introducing upstream disturbances in an inherently unstable flow.

Large-eddy simulation of a hypersonic turbulent boundary layer over a compression corner

In this paper, large-eddy simulation of the interaction between a shock wave and the hypersonic turbulent boundary layer in a compression corner with a fixed 34° deflection angle at Ma = 6 for different Reynolds number cases is conducted. For investigating the effects of the Reynolds number for hypersonic cases, three cases where the free-stream Reynolds numbers are 14000, 20000, and 30000/mm are selected. The averaged statistics, such as the mean velocity, the skin friction, the heat flux, and the wall pressure, are used in this paper. The flow structures in the compression ramp including the shock wave and interaction region are discussed. The decomposition of the mean skin-friction drag for the flat flow is extended to be used in the compression corner. In addition, the turbulent kinetic energy is studied through the decomposition of the mean skin-friction drag for the flat-plate region and the corner region. It is found that higher Reynolds numbers would increase the turbulent kinetic energy by turbulent dissipation at the interaction region, while higher Reynolds numbers would decrease the turbulent kinetic energy by turbulent dissipation after reattachment. In addition, it is also found that the turbulent kinetic energy is larger with a higher Reynolds number and higher turbulent kinetic energy inhibits the movement from the separation point to the inflection point (x = 0 mm), which deduces larger separation bubbles.

High-speed aerodynamics control using energy injection of pulsed discharges

We present a novel method of high-speed aerodynamic control that relies on active energy injection of pulsed discharges characterized by a high power density and long pulse width. The control principle, which is based on the manipulation of a shock wave by the energy injection, is investigated and verified via M5 wind tunnel experiments and a numerical simulation. The test model is a cone-shaped model with two compression ramps, and we organize four discharges which are connected in series and located on the surface of the model. We use high-resolution schlieren imaging and a fiber optic balance to analyze the control effects of these discharges on the flow topology and the aerodynamic force. Schlieren imaging reveals a notable change in the flow topology and a weakening of the attached shock wave due to the discharge plasma layer creating a virtual wedge that diverts the incoming flow and reducing the local Mach number. The quantitative balance data show a significant force reduction in the presence of the pulsed discharges, and reveals that the force reduction rate is associated with the energy ratio, which increases with an increase in the pulse energy and a decrease in the dynamic pressure. A larger energy ratio of discharge leads to a higher penetration depth of the plasma layer, which is more conducive to aerodynamic control. The simulation results show similar flow topologies and consistent variation tendency in the aerodynamic force to the experimental results.

Turbulence amplification in the shock wave/turbulent boundary layer interaction over compression ramp by the flux reconstruction method

Wall-resolved large eddy simulation on a supersonic turbulent boundary layer over a 24° compression ramp is performed under the framework of high order discontinuous methods for a free-stream Mach number M∞=2.9 and Reynolds number Re∞=5.966×106. The turbulent flow is resolved by the high order flux reconstruction method, and the shock is captured by a high-resolution, but stable weighted essentially non-oscillation limiter. The solver used in this paper is validated by the double Mach reflection case and the Taylor–Green vortex case. The results of shock wave/turbulent boundary layer interaction at the ramp corner are validated by the numerical simulations and the experimental data in the literature. The analysis of the physics behind the turbulence amplification at around the ramp corner is presented. The shear effects and the flow deceleration/acceleration are the main reasons of the turbulence amplification.

Peak heat flux prediction of hypersonic flow over compression ramp under vibrationally excited free-stream condition

In hypersonic shock tunnel experiments, the high-temperature reservoir gas expands and accelerates so rapidly that there is not enough time for vibrational energy relaxation. As a result, thermal nonequilibrium gas flow is frequently encountered in the test section, and this significantly affects the measured heat flux. In this paper, hypersonic compression-ramp flows are studied numerically to investigate the effect of incomplete vibrational energy accommodation on the separation flow structure and peak heat flux in the reattachment region under low-to-medium Reynolds number and high Mach number conditions. Numerical results and theoretical analysis suggest that the vibrational energy accommodation has no noticeable impact on the length scale of the separation zone, but strongly influences the peak heat flux of the separated ramp flows. Decomposing the peak heat flux into translational–rotational energy and vibrational energy components, qtr and qv, respectively, we find that qv/qtr characterizes the nonequilibrium degree of the vibrational energy accommodation. A formula for predicting the peak heat flux is then proposed, taking the effect of incomplete vibrational energy accommodation into consideration. Finally, surface heat flux measurements in a hypersonic shock tunnel indicate that a deviation of up to 13% in total peak heat flux could arise if vibrational energy accommodation is not considered under the vibrationally excited free-stream condition.

Wall pressure fluctuations in supersonic boundary layers over compression ramps with different turning angles

In the present study, we investigate influences of shock intensity on wall pressure fluctuations by performing direct numerical simulations of supersonic turbulence boundary layers over compression ramps with different turning angles. We found that as the turning angle increases, low-frequency motions of the separation shock are enhanced, accompanied by enlarged energetic pressure structures with lower convection velocities. By inspecting wavenumber-frequency spectra under the assumption of streamwise homogeneity, we further identified two energetic modes convected at different velocities. The one with the lower convection velocity, namely, the “slow mode,” inherited from the upstream pressure fluctuations of the turbulent boundary layer, is decelerated when passing through the oblique shock, during which the “rapid mode” with pressure fluctuations convected at higher speeds are generated. The increasing turning angle decelerates the slow mode and intensifies the fast mode. The reconstruction of the flow field suggests that the rapid mode is associated with the shear layer generated adjacent to the interaction zone, while the slow mode is associated with the Görtler vortices on the ramp.

Direct numerical simulation of compressible turbulence accelerated by graphics processing unit: An open-source high accuracy accelerated computational fluid dynamic software

This paper introduces open-source computational fluid dynamics software named open computational fluid dynamic code for scientific computation with graphics processing unit (GPU) system (OpenCFD-SCU), developed by the authors for direct numerical simulation of compressible wall-bounded turbulence. This software is based on the finite difference method and is accelerated by the use of a GPU, which provides an acceleration by a factor of more than 200 compared with central processing unit (CPU) software based on the same algorithm and a number of Message Passing Interface processes, and the running speed of OpenCFD-SCU with just 512 GPUs exceeds that of CPU software with 130 000 CPUs. GPU-Stream technology is used to implement overlap of computing and communication, achieving 98.7% parallel weak scalability with 24 576 GPUs. The software includes a variety of high-precision finite difference schemes and supports a hybrid finite difference scheme, enabling it to provide both robustness and high precision when simulating complex supersonic and hypersonic flows. When used with the wide range of supercomputers currently available, the software should be able to improve the performance of large-scale simulations by up to two orders on the computational scale. Then, OpenCFD-SCU is applied to a validation and verification case of a Mach 2.9 compression ramp with mesh numbers up to 31.2 × 109.

Direct numerical simulation of shock wave/turbulent boundary layer interaction in a swept compression ramp at Mach 6

Swept compression ramps widely exist in supersonic/hypersonic vehicles and have become a typical standard model for studying three-dimensional (3D) shock wave/turbulent boundary layer interactions (STBLIs). In this paper, we conduct a direct numerical simulation of swept compression ramp STBLI with a 34° compression angle and a 45° sweep angle at Mach 6 using a heterogeneous parallel finite difference solver. Benefitting from the powerful computing performance of the graphics processing unit, the computational grid number exceeds 5 × 106 with the spatiotemporal evolution data of hypersonic 3D STBLI obtained. The results show that the flow of the hypersonic swept compression ramp follows the quasi-conical symmetry. A supersonic crossflow with helical motion appears in the interaction region, and its velocity increases along the spanwise direction. Fluids from the high-energy-density region pass through the bow shock at the head of the main shock and crash into the wall downstream of the reattachment, resulting in the peaks in skin friction and heat flux. The peak friction and heating increase along the spanwise direction because of the spanwise variation in the shock wave inclination. In the interaction region, the unsteadiness is dominated by the mid-frequency motion, whereas the low-frequency large-scale motion is nearly absent. Two reasons for the lack of low-frequency unsteadiness are given: (1) The separation shock is significantly weaker than the reattachment shock and main shock; and (2) because of the supersonic crossflow, the perturbations propagating at the sound speed are not self-sustaining but flow along the r-direction and toward the spanwise boundary.

Numerical Investigation of the Hypersonic Inlet under Throttling with Heat Source

The influence of using a heat source to manage the shock wave boundary layer interactions (SWBLI) at the hypersonic inlet under throttling were studied numerically. This hypersonic inlet was created for a fluid flow Mach number of 5. The throttling was induced by a plug placed near the intake isolator's outlet. The study's parameters included the heat source power and size. The intake performance indicators were the total pressure recovery and the flow distortion. The position of the heat source was determined by studying the interplay of the shock waves from the compression ramp. The results demonstrated the existence of the shock waves at the heat source, and its influences on the SWBLI inside the isolator. This behaviour, led to an increase in the total pressure recovery and reduction of the flow distortion.

Investigation of the secondary flow by convergent–divergent riblets in a supersonic turbulent boundary layer over a compression ramp

In this paper, the effect of the secondary flow induced by convergent–divergent riblets in supersonic turbulent boundary layers over a 24° compression ramp at Mach number 2.9 is studied via direct numerical simulation. Two riblet cases with the wavelength Λ being 1.1δ and 1.65δ (δ is the boundary layer thickness) are conducted to examine their impact on the secondary rolling motion, momentum transfer, turbulent fluctuations, flow separation, and unsteady shock motion. As the flow develops over the riblet section, both the size and intensity of the secondary rolling motion tend to increase. For the riblet case with Λ/δ=1.1, a single rolling mode is observed within a half wavelength, while a pair of co-rotating vortical structures is obtained for Λ/δ=1.65. Both rolling patterns lead to an apparent spanwise variation of the flow field. The results reveal that the secondary flow contributes to the increase of both the mean momentum flux and turbulent fluctuations. By using the spanwise averaging, the mean momentum flux contributed from the dispersive stress and compressible effect caused by the secondary flow is identified. Both components appear to enhance the near-wall momentum mixing, and a larger enhancement is observed for Λ/δ=1.1, where the intensity of the secondary flow is stronger. Compared to the baseline case, the area of the separation zone at Λ/δ=1.1 and Λ/δ=1.65 is decreased by 56% and 38%, respectively. For all the cases, the low-frequency motion near the foot of the shock is observed. In comparison, the frequency of the low-frequency motion for the riblet case is two times higher than that in the baseline case, owing to the reduction of the separation area and length.

A study on flow control in a hypersonic inlet using a plasma synthetic jet actuator

The plasma synthetic jet actuator (PSJA) has been proven to have potential for shock wave control in the supersonic/hypersonic flow. To evaluate the possibility for the PSJA applied in the hypersonic inlet under a Mach number greater than the design value and provide a deeper understanding of its control mechanism, the relevant two-dimensional numerical simulations are carried out. Based on the discussions about the numerical results, the effects are clear of the actuations generated by the PSJAs, including diffracted waves, synthetic jets, and discharged heat, on the oblique shock waves induced by the compression ramps and the aerodynamic performance of the hypersonic inlet. Concretely, the shapes of the oblique waves are changed by propagations of the diffracted waves, whose starting point moved due to the synthetic jets. A deceleration occurs near the surface of the hypersonic inlet caused by the movement of the discharged heat, leading to a thickener boundary layer. Furthermore, the incremental analysis about the total pressure recovery coefficient at the throat shows that the diffracted waves can increase the static pressure at the throat achieving a higher total pressure recovery coefficient, while the discharged heat makes the total pressure recovery coefficient drop due to a lower local Mach number at the throat. The effect of the discharged heat is more obvious than that of the diffracted waves.

An immersed boundary method for wall-modeled large-eddy simulation of turbulent high-Mach-number flows

The excessive computational resource requirements for direct numerical simulation (DNS) and wall-resolved large eddy simulation (LES) make these methods intractable for simulations concerning turbulent flows over complex, real-world geometries. Wall-modeled LES (WMLES) approaches seek to reduce this cost by applying wall-layer models to the inner-most region of turbulent boundary layers. Immersed boundary methods (IBMs) have grown in popularity over the past few decades since the mesh can be automatically generated for any complex geometry - a potentially major advancement for common industrial CFD simulations. However, immersed boundary approaches suffer from reduced numerical accuracy and, possibly, stability at the wall, primarily attributable to irregular discretization of boundary operators. This is a problem that is especially significant for WMLES of high Mach number flows, where numerical schemes have the conflicting requirements to be both stable at high wavenumbers and non-dissipative. This presentation introduces a coupled IBM-WMLES approach that is tailored to the simulation of high-speed turbulent flows. The basis of this method is to carefully address numerical dissipation and stability away from and close to the wall. This is achieved by using a hybrid central-upwind flux reconstruction scheme and careful application of smooth boundary conditions. The presented test cases evaluate the current scheme for a wide range of physical flow phenomena, and include a turbulent channel for basic validation, a hypersonic transitional boundary layer to investigate high-speed zero-pressure-gradient flows, a hypersonic shock wave boundary layer interaction to investigate adverse pressure gradients, separation, and turbulent heating, and a hypersonic compression ramp which is significantly challenging as it combines all of the features from the aforementioned high-speed flow problems.

Assessing the Performance of Hypersonic Inlets by Applying a Heat Source with the Throttling Effect

Utilization of a heat source to regulate the shock wave–boundary layer interaction (SWBLI) of hypersonic inlets during throttling was computationally investigated. A plug was installed at the intake isolator’s exit, which caused throttling. The location of the heat source was established by analysing the interaction of the shockwave from the compression ramp and the contact spot of the shockwave with that of the inlet cowl. Shockwave interaction inside the isolator was investigated using steady and transient cases. The present computational work was validated using previous experimental work. The flow distortion (FD) and total pressure recovery (TPR) of the inflows were also studied. We found that varying the size and power of the heat source influenced the shockwaves that originated around it and affected the SWBLI within the isolator. This influenced most of the performance measures. As a result, the TPR increased and the FD decreased when the heat source was applied. Thus, the use of a heat source for flow control was found to influence the performance of hypersonic intakes.

Influence of jet spacing in spanwise-inclined jet injection in supersonic crossflow

In separation control with air-jet vortex generators in supersonic flow, the spacing of the jets in a row array has a crucial effect on the control effectiveness. Previous experimental studies have revealed the overall influence of jet spacing on air-jet vortex generator controlled shock-induced separation zones, focusing mostly on the separation regions. There is, however, a gap in knowledge regarding the mechanisms leading to these changes in control effectiveness, particularly on the interaction between multiple jets and the details of downstream flow evolution. Therefore, the objective of the current study is to provide detailed information on the underlying flow dynamics associated with the injection of a row of spanwise-inclined jets into a supersonic turbulent boundary layer – and in particular the effects of different jet spacings. Four different spacings were studied with large-eddy simulations. In addition, we performed oil-flow and schlieren visualizations of separation control in a 24$^\circ$compression–ramp interaction with different jet-spacing configurations to validate and discuss our conclusions regarding the effects of jet/jet interactions on the separation-control effectiveness. We analyse the influence of jet spacing on the flow topology, induced vortical structures, and boundary-layer statistics. The jet-induced major vortex pair is the dominant flow structure energizing the near-wall boundary layer. The paper details how interactions amongst adjacent vortex pairs differ for varying jet spacings, thus influencing the momentum transfer and eventually the control efficiency. The dynamic behaviour of the flow was analysed using a three-dimensional dynamic mode decomposition. The resulting insights are key to the development of efficient control set-ups.

Direct numerical simulation of shock-wave/boundary layer interaction controlled with convergent–divergent riblets

In this paper, a section of convergent–divergent (C–D) riblets is applied upstream of a compression ramp in a supersonic turbulent boundary layer in a Mach 2.9 flow with a Reynolds number of Reθ=2240. Direct numerical simulations are undertaken to examine the impact of C–D riblets on the shock wave/boundary layer interaction and the feasibility of using them to mitigate flow separation. Over the riblet section, a large-scale secondary roll mode is produced by C–D riblets with the downwelling motion occurring around the diverging region and upwelling motion near the converging region. This consequently leads to a spanwise heterogeneity in mean quantities and turbulent structures over the riblet section and also in the interaction zone. Compared with the baseline case, the area of the separation zone for the riblet case experiences a dramatic local reduction of 92% in the diverging region, owing to the downwelling motion that injects the high-momentum fluid toward the wall and the near-wall spanwise velocity that transports the low-momentum fluid away. The enhanced upwelling motion around the converging region induced by C–D riblets, on the one hand, contributes to the decrease of the near-wall momentum and subsequently the increase of the local separation area. On the other hand, the upwelling motion effectively reduces the incoming Mach number upstream of the compression corner. This appears to reduce the strength of the separation shock, leading to a more gradual compression of the incoming flow that helps ease the enlargement of the separation area nearby. Overall, the area of the mean flow separation is reduced by 56%, indicating an effective flow separation control by the C–D riblets.

Comminuted intraarticular calcaneal fractures: Multiplanar VA locked plating and interlocked nailing incorporate longitudinal strut and provide superior stability – a biomechanical cadaveric study

Treatment of comminuted intraarticular calcaneal fractures remains controversial and challenging. The aim of this study was to investigate the biomechanical performance of three different methods for fixation of such fractures. Comminuted calcaneal fractures, including Sanders III AB fracture of the posterior facet and Kinner II B fracture of the calcaneocuboid joint (CCJ) articular calcaneal surface, were created in 18 human cadaveric lower legs by osteotomizing. The ankle joint, medial soft tissues and midtarsal bones along with their ligaments were preserved. The specimens were randomized to three groups for fixation with either (1) 2.7 mm variable-angle locking lateral calcaneal plate (Group 1), (2) 2.7 mm variable-angle locking anterolateral calcaneal plate in combination with one 4.5 mm and one 6.5 mm cannulated screws (Group 2), or (3) interlocking calcaneal nail with 3.5 mm screws in combination with three separate 4.0 mm cannulated screws (Group 3). All specimens were biomechanically tested to failure under axial loading in midstance foot position. Each test commenced with a quasi-static compression ramp from 50 to 200 N, followed by progressively increasing cyclic loading at 2 Hz. Starting from 200 N, the peak load of each cycle increased at a rate of 0.2 N/cycle. Interfragmentary movements were captured by motion tracking. In addition, mediolateral X-rays were taken every 250 cycles with a triggered C-arm. Böhler angle after 5000 cycles (1200 N peak load) increased significantly more in Group 1 compared to both other groups (P ≤ 0.020). Varus deformation of 10° between the calcaneal tuberosity and the lateral calcaneal fragments was reached at significantly lower number of cycles in Group 1 compared the other groups (P ≤ 0.017). Both cycles to 10° plantar gapping between the anterior process and the calcaneal tuberosity fragments, and 2 mm displacement at the CCJ articular calcaneal surface revealed no significant differences among the groups (P ≥ 0.773). From a biomechanical perspective, treatment of comminuted intraarticular calcaneal fractures using anterolateral variable-angle locking plate with additional longitudinal screws or interlocked nail in combination with separate transversal screws provides superior stability as opposed to lateral variable-angle locked plating only.

Control effects of a high-frequency pulsed discharge on a hypersonic separated flow

The control effects of a high-frequency pulsed discharge on a hypersonic separated flow induced by a 42° compression ramp are studied with high-speed schlieren imaging. A discharge is generated between an electrode pair upstream of the ramp to introduce strong disturbances into the boundary layer. The discharge frequencies used are fd = 10, 20, 30, and 40 kHz, with corresponding single-pulse energies Esp = 22.2, 20.3, 12.0, and 8.7 mJ. The results show that the discharge can reduce the separation area, especially with relatively low frequencies of 10 or 20 kHz. The flow mechanism induced by the discharge is scrutinized using a reduced-order method based on the information from time-resolved schlieren images, including proper orthogonal decomposition and dynamic mode decomposition. The results identify an unsteady mode representing regular energetic eddies within the shear layer of the separation zone. This mode has the same characteristic frequency as the discharge. These eddies enhance the mixing of the shear layer and the main flow stream, accelerating the destruction of the shear layer and reducing the separation zone size. The control effectiveness correlates with the scale of the energetic eddies. At frequencies of 10 or 20 kHz, the discharge is more conducive to attenuating the separation because large-scale eddies, excited by higher single-pulse energy, are enhanced as they propagate downstream in the shear layer. By contrast, the control effectiveness deteriorates when the discharge frequency is 30 or 40 kHz because the induced eddies have smaller scales and decay during propagation.

Plasma-based experimental investigation of double compression ramp shock wave/boundary layer interaction control

An experimental investigation on the control effects of the high-frequency streamwise pulsed arc discharge array (HS-PADA) on the double compression ramp shock wave/boundary layer interaction (DCR-SWBLI) was carried out at Mach 2.0. Firstly, two types of ramp configurations were designed. The base flow field and actuation flow field were investigated. The actuation frequencies were 10 kHz and 20 kHz. Fast Fourier transform and root-mean-square methods were applied based on schlieren images. The base flow field indicated that the compression effect of ramps decreased with the lengthen of the first ramp. The results of actuation flow field showed that the 20 kHz actuation was superior to the 10 kHz actuation in weakening the shock wave intensity. The HS-PADA exhibited two types of control effects: modifying the shock structure, which was closely related to the separation zone, and modifying the low-frequency unsteadiness of the shock wave, which might not be related to the separation zone. The first separation shock wave, of which the high-frequency motion was more intense under actuation, may be more sensitive to HS-PADA than the second. Finally, the control mechanism of the HS-PADA on DCR-SWBLI was extracted.