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A special issue of Aerospace (ISSN 2226-4310). This special issue belongs to the section "Aeronautics".

Deadline for manuscript submissions: closed (31 March 2024) | Viewed by 9326

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Dr. Hexia Huang

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Guest Editor

College of Energy and Power Engineering, Nanjing University of Aeronautics and Astronautics, Nanjing 210016, China
Interests: aerodynamics in high-speed inlet; internal flow; shock/boundary layer interaction; flow control
Special Issues, Collections and Topics in MDPI journals

Dr. Ye Tian

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Guest Editor

China Aerodynamic Research and Development Center, Mianyang 621000, China
Interests: hypersonic aerodynamics; scramjet combustion; high-speed propulsion; experimental fluid dynamics; shock train

Prof. Dr. Huijun Tan

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Guest Editor

Special Issue Information

Dear Colleagues,

The shock-dominated flow is frequently encountered in high-speed aircraft and engines, and its flow characteristics directly determine the aerodynamic performance of the aircraft and engines. Due to the strong discontinuity and pressurization property of shock waves, the shock-dominant flow exhibits strong non-linearity, strong inviscid/viscous interaction, and significant historical effects, making it difficult to predict the related flow structures and behaviors. With the development of aircraft towards higher speeds, better performance, and more intelligent control, the shock-dominated flow is a key scientific issue, involving complex high-speed aerodynamics, flow stability, fluid/ thermal structure/acoustic multi-fields interaction, flow control, and artificial intelligence. For this reason, this Special Issue focuses on the shock/boundary layer interaction, shock/shock interaction, shock/vortex interaction, shock oscillation, shock-dominated flow in high-speed aircraft/engines, CFD and experimental methods in shock-dominated, shock aerodynamic heating and cooling, shock/thermal structure interaction, etc.

The editor of this Special Issue invites authors to submit papers on addressing the challenges in the shock-dominated flow of various aerospace occasions.

Dr. Hexia Huang
Dr. Ye Tian
Prof. Dr. Huijun Tan
Guest Editors

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Keywords

shock/boundary layer interaction
shock/shock interaction
shock oscillation
shock wave flow in high-speed engine
computational methods in shock dominated flow
experimental methods in shock dominated flow
shock wave dominated flow control
shock heating
shock dominated flow cooling
shock wave/thermal structure interaction
explosion shock
turbulent combustion with shocks
application of shock wave in the engineering field

Published Papers (9 papers)

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Research

25 pages, 14770 KiB

Open AccessArticle

Experimental Investigation of the Shock-Related Unsteadiness around a Spiked-Blunt Body Based on a Novel DMD Energy Sorting Criterion

by Yifan Wang, Jinglei Xu, Qihao Qin, Ruiqing Guan and Le Cai

Aerospace 2024, 11(3), 188; https://doi.org/10.3390/aerospace11030188 - 27 Feb 2024

Viewed by 855

Abstract

In this study, we propose a novel dynamic mode decomposition (DMD) energy sorting criterion that works in conjunction with the conventional DMD amplitude-frequency sorting criterion on the high-dimensional schlieren dataset of the unsteady flow of a spiked-blunt body at Ma = 2.2. The study commences by conducting a comparative analysis of the eigenvalues, temporal coefficients, and spatial structures derived from the three sorting criteria. Then, the proper orthogonal decomposition (POD) and dynamic pressure signals are utilised as supplementary resources to explore their effectiveness in capturing spectral characteristics and spatial structures. The study concludes by summarising the characteristics and potential applications of DMD associated with each sorting criterion, as well as revealing the predominant flow features of the unsteady flow field around the spiked-blunt body at supersonic speeds. Results indicate that DMD using the energy sorting criterion outperforms the amplitude and frequency sorting criteria in identifying the primary structures of unsteady pulsations in the flow field, which proves its superiority in handling an experimental dataset of unsteady flow fields. Moreover, the unsteady pulsations in the flow field around the spiked-blunt body under supersonic inflow conditions are observed to exhibit multi-frequency coupling, with the primary frequency of 3.3 kHz originating from the periodic motion of the aftershock. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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16 pages, 8093 KiB

Open AccessArticle

Mitigation of Shock-Induced Separation Using Square-Shaped Micro-Serrations—A Preliminary Study

by Fangyou Yu, Zhanbiao Gao, Qifan Zhang, Lianjie Yue and Hao Chen

Aerospace 2024, 11(2), 148; https://doi.org/10.3390/aerospace11020148 - 12 Feb 2024

Viewed by 809

Abstract

Suppressing shock-induced flow separation has been a long-standing problem in the design of supersonic vehicles. To reduce the structural and design complexity of control devices, a passive control technique based on micro-serrations is proposed and its controlling effects are preliminarily investigated under test conditions in which the Mach number is 2.5 and the ramp creating an incident shock is 15 deg. Meanwhile, a vorticity-based criterion for assessing separation scales is developed to resolve the inapplicability of the zero skin friction criterion caused by wall unevenness. The simulations demonstrate that the height of the first stair significantly influences the separation length. Generally, the separation length is shorter at higher stairs, but when the height is greater than half of the thickness of the incoming boundary layer, the corresponding separation point moves upstream. A stair with a height of only 0.4 times the thickness of the boundary layer reduces the separation length by 2.69%. Further parametric analysis reveals that while the remaining serrations have limited effects on the flow separation, an optimization of their shape (depth and width) can create more favorable spanwise vortices and offer a modest improvement of the overall controlling performance. Compared to the plate case, a 9.13% reduction in the separation length can be achieved using a slightly serrated design in which the leading stair is 0.1 high and the subsequent serrations are 0.2 deep and 0.05 wide (nondimensionalized, with the thickness of the incoming boundary layer). Meanwhile, the micro-serration structure even brings less drag. Considering the minor modification to the structure, the proposed method has the potential for use in conjunction with other techniques to exert enhanced control on separations. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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14 pages, 6086 KiB

Open AccessArticle

Experimental Study on Hypersonic Double-Wedge Induced Flow Based on Plasma Active Actuation Array

by Bo Yang, Hesen Yang, Ning Zhao, Hua Liang, Zhi Su and Dongsheng Zhang

Aerospace 2024, 11(1), 60; https://doi.org/10.3390/aerospace11010060 - 09 Jan 2024

Viewed by 949

Abstract

The double-wedge configuration is a typical characteristic shape of the rudder surface of high-speed aircraft. The impact of the shock wave/boundary layer interaction and the shock wave/shock wave interaction resulting from the double wedge on aircraft aerodynamics cannot be ignored. The aerodynamic performance of the aircraft would be seriously affected. Accordingly, to reduce the wave drag, and to relieve the thermal load and pressure load, flow control is required for the shock wave/shock wave interaction and the shock wave/boundary layer interaction induced by the double-wedge configuration. In this paper, double-wedge shock wave/shock wave interaction is controlled by a high-energy surface arc discharge array and observed by high-speed schlieren flow field measurement at Mach 8. The 30-channel discharge array is set on the primary wedge plane, and actuation is generated. Hypersonic V shock wave/shock wave interaction is effectively controlled by the shock wave array induced by the high-energy surface arc discharge array, which makes the shock wave/shock wave interaction structure disappear or intermittent. The potential control mechanism is to reduce strong shock wave interaction by transforming the type of shock wave interaction. Therefore, the ability of plasma array actuation to control complex shock wave/shock wave interaction is verified, which provides a new method for hypersonic shock wave/shock wave interaction control. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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20 pages, 17494 KiB

Open AccessArticle

Transient Flow Evolution of a Hypersonic Inlet/Isolator with Incoming Windshear

by Simin Gao, Hexia Huang, Yupeng Meng, Huijun Tan, Mengying Liu and Kun Guo

Aerospace 2023, 10(12), 1021; https://doi.org/10.3390/aerospace10121021 - 09 Dec 2023

Viewed by 931

Abstract

In this paper, a novel flow perturbation model meant to investigate the effects of incoming wind shear on a hypersonic inlet/isolator is presented. This research focuses on the transient shock/boundary layer interaction and shock train flow evolution in a hypersonic inlet/isolator with an on-design Mach number of 6.0 under incoming wind shear at high altitudes, precisely at an altitude of 30 km with a magnitude speed of 80 m/s. Despite the low intensity of wind shear at high altitudes, the results reveal that wind shear significantly disrupts the inlet/isolator flowfield, affecting the shock wave/boundary layer interaction in the unthrottled state, which drives the separation bubble at the throat to move downstream and then upstream. Moreover, the flowfield behaves as a hysteresis phenomenon under the effect of wind shear, and the total pressure recovery coefficients at the throat and exit of the inlet/isolator increase by approximately 10% to 12%. Furthermore, this research focuses on investigating the impact of wind shear on the behavior of the shock train. Once the inlet/isolator is in a throttled state, wind shear severely impacts the motion of the shock train. When the downstream backpressure is 135 times the incoming pressure (p₀), the shock train first moves upstream and gradually couples with a cowl shock wave/boundary layer interaction, resulting in a more significant separation at the throat, and then moves downstream and decouples from the separation bubble at the throat. However, if the downstream backpressure increases to 140 p₀, the shock train enlarges the separation bubble, forcing the inlet/isolator to fall into the unstart state, and it cannot be restarted. These findings emphasize the need to consider wind shear effects in the design and operation of hypersonic inlet/isolator. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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17 pages, 9180 KiB

Open AccessArticle

Experimental Investigation on the Control of Hypersonic Shock Wave/Boundary Layer Interaction Using Surface Arc Plasma Actuators at Double Compression Corner

by Bo Yang, Hesen Yang, Chuanbiao Zhang, Ning Zhao, Hua Liang and Dongsheng Zhang

Aerospace 2023, 10(12), 1016; https://doi.org/10.3390/aerospace10121016 - 06 Dec 2023

Viewed by 1056

Abstract

Compression corner shock wave/boundary layer interaction (SWBLI) is a typical shock wave/boundary layer interaction (SWBLI) problem in supersonic/hypersonic flows. In previous studies, the separation flow is usually caused by a single shock wave. However, in the actual aircraft surface configuration, two-stage compression or even multistage compression will produce more complex SWBLI problems. The multi-channel shock structure makes the flow field structure more complicated and also puts forward higher requirements for the flow control scheme. In order to explore a flow control method for the double compression corner shock wave/boundary layer interaction problem, an experimental study is carried out to control the double compression corner shock wave/boundary layer interaction with a high-energy flow pulsed arc discharge array under the condition that the incoming flow velocity Ma 6.0 has both noise flow fields and quiet flow fields. The results show that when U_DC = 0.5 kV actuation is applied, the influence range of the hot gas mass flow direction is about 65 mm, which can weaken the shock wave intensity to a certain extent. When U_DC = 1 kV actuation is applied, the influence range of the hot gas mass flow direction extends to 85 mm, and the actuation has a significant control effect on the flow field. Through spatio-temporal evolution analysis and spatial gradient threshold processing of high-speed schlieren images of actuated flow fields, the feasibility of controlling the hypersonic double compression corner shock wave/boundary layer interaction by using a high-energy flow pulsed arc discharge array is verified. The control law of a high-energy flow pulsed arc discharge array acting on the double compression corner shock wave/boundary layer interaction is revealed. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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20 pages, 8788 KiB

Open AccessArticle

The Influence of External Flow Field on the Flow Separation of Overexpanded Single-Expansion Ramp Nozzle

by Yang Yu, Yuepeng Mao, Tao Yu, Yalin Yang, Shulin Xu and Sijia Liang

Aerospace 2023, 10(11), 958; https://doi.org/10.3390/aerospace10110958 - 13 Nov 2023

Viewed by 1045

Abstract

Flow separation and transitions of separation patterns are common phenomena of nozzles working with a wide Mach range. The maximum thrust method is applied to design the single-expansion ramp nozzle (SERN) for specific operating conditions. The nozzle is used to numerically simulate the transition processes of separation patterns under the linear change in the external flow Mach number and the actual trajectory take-off condition of a rocket-based combined cycle (RBCC), to investigate the mechanism through which the external flow field influences the separation pattern transition during acceleration. The computational fluid dynamics (CFD) method is briefly introduced, followed by experimental validation. Then, the design procedure of SERN is described in detail. The simulation results indicate that as the external Mach number increases, the flow field in the nozzle undergoes transitions from RSS (ramp) to FSS, and finally exhibits a no-flow separation pattern. The rate at which the external Mach number varies has little effect on the transition principle of the nozzle flow separation patterns, but it has a significant effect on the critical Mach number of the transition points. The external flow field of the nozzle has an airflow accumulation effect during acceleration, which can delay the transition of the flow separation pattern. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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22 pages, 11423 KiB

Open AccessArticle

Noise Prediction and Plasma-Based Control of Cavity Flows at a High Mach Number

by Hongming Cai, Zhuoran Zhang, Ziqi Li and Hongda Li

Aerospace 2023, 10(11), 922; https://doi.org/10.3390/aerospace10110922 - 29 Oct 2023

Viewed by 923

Abstract

Cavity flows are a prevalent phenomenon in aerospace engineering, known for their intricate structures and substantial pressure fluctuations arising from interactions among vortices. The primary objective of this research is to predict noise levels in high-speed cavity flows at Mach 4 for a rectangular cavity characterized by an aspect ratio of L/D = 7. Moreover, this study delves into the influence of the plasma actuator on noise control within the cavity flow regime. To comprehensively analyze acoustic characteristics and explore effective noise reduction strategies, a computational fluid dynamics technique with the combination of a delayed detached eddy simulation (DDES) and plasma phenomenological model is established. Remarkably, the calculated overall sound pressure level (OASPL) and plasma-induced velocity closely align with the experimental data, validating the reliability of the proposed approach. The results show that the dielectric barrier discharge (DBD) plasma actuator changes the movement range of a dominating vortex in the cavity to affect the OASPL at the point with the maximum noise level. The control of excitation voltage can reduce the cavity noise by 2.27 dB at most, while control of the excitation frequency can only reduce the cavity noise by 0.336 dB at most. Additionally, the increase in excitation frequency may result in high-frequency sound pressure, but the influence is weakened with the increase in the excitation frequency. The findings highlight the potential of the plasma actuator in reducing high-Mach-number cavity noise. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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16 pages, 10299 KiB

Open AccessArticle

Evolution of Shock Waves during Muzzle Jet Impinging Moving Bodies under Different Constrained Boundaries

by Zijie Li and Hao Wang

Aerospace 2023, 10(11), 908; https://doi.org/10.3390/aerospace10110908 - 25 Oct 2023

Cited by 1 | Viewed by 882

Abstract

A recently developed launching device called the gun–track launch system is affected by its constrained track, such that the form of the muzzle jet changes from the state of free development in the entire space to a constrained state, where this lends unique characteristics of development to its flow field. In this study, the authors establish the corresponding model for numerical simulations based on the dynamic mesh method. We also considered a model of simulation of the muzzle jet with an “infinitely” constrained track to analyze its performance under real launch conditions to explore the mechanism of development and the disturbance-induced propagation of the shock wave when the muzzle jet impinges on moving bodies. The results showed that the muzzle jet exhibited a circumferential asymmetric shape that tilted toward the area above the muzzle and generated transverse air flow that led to the generation of a vortex on it. Because the muzzle was close to the ground, the jet was reflected by it to enhance the development and evolution of the shock waves and vortices and to aggravate the rate of distortion and asymmetry of the jet. The wave reflected from the ground was emitted once again when it encountered the infinitely constrained track. No local low-pressure area or a prominent vortex was observed after multiple reflections. Because the track in the test model was short, the waves reflected by the ground were not blocked, and vortices were formed in the area above the ground. Significant differences in the changes in pressure were also observed at key points in the domain. The results of a comparative analysis showed that the infinitely constrained track increased the Mach number of the moving body from 1.4 to 1.6. The work provides a theoretical basis and the requisite technical support for applications of the gun–track launch system. Full article

(This article belongs to the Special Issue Shock-Dominated Flow)

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26 pages, 12223 KiB

Open AccessArticle

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

by Mengge Wang, Ziyun Wang, Yue Zhang, Daishu Cheng, Huijun Tan, Kun Wang and Simin Gao

Aerospace 2023, 10(8), 729; https://doi.org/10.3390/aerospace10080729 - 20 Aug 2023

Cited by 1 | Viewed by 1176

Abstract

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

(This article belongs to the Special Issue Shock-Dominated Flow)

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