Structural characteristics of the shock-induced boundary layer separation extended to the leading edge

Tao, Yuan; Liu, W. D.; Fan, Xiaoqiang; Zhao, Yunfei

doi:10.1063/1.4993756

Cited by 16 publications

(3 citation statements)

References 20 publications

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“…In overrated status, the shocks and compression waves produced by the forebody should have gathered in the leading edge of the lip impact the inner surface of the lip. Then, large-scale separated flow is induced, and a suspended shock structure typically appears near the leading edge of the lip [36,37]. Jiao and Wang conducted systematic research into overrated unstart and clarified the flow characteristics and variation in the regularity of overrated unstart.…”

Section: Introductionmentioning

confidence: 99%

Self-Start Characteristics of Hypersonic Inlet When Multiple Unstart Modes Exist

Tang,

Xiong,

Fan

et al. 2023

Applied Sciences

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Intense shock boundary-layer interaction may lead to multiple unstart modes existing in a hypersonic inlet. Thus, self-start problems become complex and cannot be explained using the classical double-solution theory of air inlet. The essence of the self-start process of a hypersonic inlet is the vanishment of separations near or in the inlet. To clarify self-start characteristics, experiments were conducted on three distinct types of unstart mode: the flow mode of small separation on body (SSB), large separation on body (LSB), and dual separations on both body and lip (DSBL); researchers recently discovered these as the unstart modes of hypersonic inlet. The results from the current experiment are as follows: (1) The SSB vanishes by raising the angle of attack (alpha). Before the vanishing point is reached, there is a dwindling process for this separation. (2) The LSB vanishes through acceleration or a decreasing alpha. (3) DSBL are difficult to vanish directly, which results in poor self-start performance. However, the DSBL flow mode may convert to LSB unstart form—which is easier to self-start—by decreasing the alpha. The Flow Field Reconstruction Method was designed to improve the self-start of the DSBL flow mode, and it was validated through experiments. Analysis of the flow mechanism revealed the reason for the poor self-start performance of the DSBL unstart mode: large-scale separation on the lip side cannot be promoted to vanish through broadwise spillage due to the resistance of sideboards. The results of this study could greatly enrich the existing theory of start problems for hypersonic inlets.

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Section: Introductionmentioning

confidence: 99%

Self-Start Characteristics of Hypersonic Inlet When Multiple Unstart Modes Exist

Tang,

Xiong,

Fan

et al. 2023

Applied Sciences

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“…It can highly affect the performance of a propulsion system. SWBLI generates an adverse pressure gradient at the interaction region with the consequences of flow separation, increased viscous dissipation, increased drag, and increased thickness of boundary layer [2]. Figure 1 shows a schematic of the SWBLI structure in a supersonic flow over a shock generator including incident and separation shocks, reflected shock, sonic line, and separation point [3].…”

Section: Introductionmentioning

confidence: 99%

Investigation on Supersonic Flow Control Using Nanosecond Dielectric Barrier Discharge Plasma Actuators

Shahrbabaki

Bazazzadeh

Khoshkhoo

2021

International Journal of Aerospace Engineering

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In this paper, the effects of streamwise Nanosecond Dielectric Barrier Discharge (NS-DBD) actuators on Shock Wave/Boundary Layer Interaction (SWBLI) are investigated in a Mach 2.5 supersonic flow. In this regard, the numerical investigation of NS-DBD plasma actuator effects on unsteady supersonic flow passing a 14° shock wave generator is performed using simulation of Navier-Stokes equations for 3D-flow, unsteady, compressible, and k ‐ ω SST turbulent model. In order to evaluate plasma discharge capabilities, the effects of plasma discharge length on the flow behavior are studied by investigating the flow friction factor, the region of separation bubble formation, velocity, and temperature distribution fields in the SWBLI region. The numerical results showed that plasma discharge increased the temperature of the discharge region and boundary layer temperature in the vicinity of flow separation and consequently reduced the Mach number in the plasma discharge region. Plasma excitation to the separation bubbles shifted the separation region to the upstream around 6 mm, increased SWBLI height, and increased the angle of the separation shock wave. Besides, the investigations on the variations of pressure recovery coefficient illustrated that plasma discharge to the separation bubbles had no impressive effect and decreased pressure recovery coefficient. The numerical results showed that although the NS-DBD plasma actuator was not effective in reducing the separation area in SWBLI, they were capable of shifting the separation shock position upstream. This feature can be used to modify the structure of the shock wave in supersonic intakes in off-design conditions.

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“…Shock-wave/boundary-layer interactions (SWBLI) have significant influences on supersonic vehicles because they cause aerodynamic issues such as increased drag, high heat loading, and adverse pressure gradients 1 , hence, the physics of SWBLI has been widely investigated both experimentally and numerically 2,3 . In isolators of high-speed airbreathing engines for example, the adverse pressure gradient caused by SWBLI sometimes induces flow separation, which can lead to a reverse flow in the boundary layer, and a decrease in total pressure resulting in significant performance loss or unstart of these engines.…”

Section: Introductionmentioning

confidence: 99%

Numerical investigation of nanosecond pulsed plasma actuators for control of shock-wave/boundary-layer separation

2018

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This study numerically explores the flow physics associated with nanosecond pulsed plasma actuators that are designed to control shock-wave induced boundary-layer separation in a Mach 2.8 supersonic flow. By using two dielectric barrier surface discharge actuator configurations, parallel and canted with respect to the flow velocity vector, a previous experiment suggested that the actuator worked in two ways to influence the interaction: boundary layer heating and vorticity production. The heating effect was enhanced with the parallel electrode and made the boundary-layer separation stronger, while the canted electrode produced vorticity and suppressed the boundary-layer separation due to the momentum transfer from the core flow. Because the detailed physical processes are still unclear, in this paper a numerical investigation is undertaken with a large eddy simulation and an energy deposition model for the plasma actuation, in which the dielectric barrier discharge produced plasma is approximated as a high temperature region. The flow characteristics without the plasma actuation correspond to the experimental observation, indicating that the numerical method successfully resolves the shock-wave/boundary-layer interaction. With the plasma actuation, complete agreement between the experiment and calculation has not been obtained in the size of shock-wave/boundary-layer interaction region. Nevertheless, as with the experiment, the calculation successfully demonstrates definite difference between the parallel and canted electrodes: the parallel electrode causes excess heating and increases the strength of the interaction, while the canted electrode leads to a reduction of the interaction strength, with a corresponding thinning of the boundary layer due to the momentum transfer. The counter flow created by the canted actuator plays an important role in the vortex generation, transferring momentum to the boundary layer and, consequently, mitigating the shock induced boundary layer separation.

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Structural characteristics of the shock-induced boundary layer separation extended to the leading edge

Cited by 16 publications

References 20 publications

Self-Start Characteristics of Hypersonic Inlet When Multiple Unstart Modes Exist

Self-Start Characteristics of Hypersonic Inlet When Multiple Unstart Modes Exist

Investigation on Supersonic Flow Control Using Nanosecond Dielectric Barrier Discharge Plasma Actuators

Numerical investigation of nanosecond pulsed plasma actuators for control of shock-wave/boundary-layer separation

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