This study investigates the effects of Supersonic Retropropulsion (SRP) on flow structures and aerodynamic loads during the vertical landing of reusable launch vehicles. A combined approach utilizing hypersonic wind tunnel experiments and Reynolds-averaged Navier-Stokes simulations is employed to systematically analyze the evolution of SRP-induced flow interference under varying freestream and thrust conditions. The investigation focuses on the complex interaction mechanisms between the retropropulsive jet and the hypersonic freestream, the evolution characteristics of vortex ring structures and their modulation of flow unsteadiness, as well as the quantitative relationships between characteristic flow scales and thrust parameters. The results show that SRP significantly alters the aerodynamic loading and flowfield topology in the aft-body region. The introduction of angle of attack induces asymmetric vortex ring evolution and large-scale low-frequency disturbances, which further affect the dynamic stability of the triple-point, shock structures, and recirculation zones. Under asymmetric conditions, the angle between the vortex core line and the wind tunnel axis rapidly decreases from 90° to 72.13° within 300 μs, corresponding to a transverse expansion velocity of 73.3 m/s. These asymmetric dynamics exhibit markedly stronger evolution rates and flow modulation compared to symmetric cases. Both experimental and numerical results consistently show a linear relationship between the locations of key flow structures and the square root of the thrust coefficient (CT). Additionally, the maximum Mach number increases almost linearly from 5.77 to 6.73 with increasing CT, showing minimal sensitivity to angle of attack. These findings enhance the physical understanding of SRP flow interference and offer valuable guidance for the development of predictive aerodynamic models and control strategies.

Surface Pulsed Arc Discharge (SPAD) plasma actuators have demonstrated significant potential for active flow control in high-speed flow environments. The paper studies the impact of SPAD on the aerodynamic performance of hypersonic airfoils in a Mach 5 flow field through numerical simulations. The results reveal that both the geometric placement of the actuator and its discharge parameters have a critical influence on the actuation efficacy in enhancing aerodynamic performance. Specifically, the aerodynamic control effect of SPAD is primarily manifested in two complementary aspects: lift enhancement and drag reduction. Notably, actuator configurations dominated by lift increase induce substantial alterations in the airfoil’s pitching moment. When the SPAD is deployed within the airfoil’s transitional region, peak improvements are achieved: a maximum lift-to-drag ratio enhancement of up to 20.7% and a corresponding maximum aerodynamic performance change of 24.9%. Under a constraint of fixed total electrical energy input, optimal aerodynamic outcomes can be achieved by tuning the discharge frequency and pulse width. At low discharge frequencies, the flow structures generated by individual pulses exhibit minimal temporal interference, allowing their aerodynamic effects to accumulate nearly linearly over successive pulses. This quasi-independent pulse interaction enables a time-additive control strategy that effectively enhances overall performance. The findings provide valuable insights and practical guidelines for designing and implementing SPAD-based flow control systems in hypersonic applications, highlighting its promise as a versatile and energy-efficient actuation technology for next-generation high-speed vehicles.

Dielectric Barrier Discharge(DBD)plasma actuators have demonstrated significant potential for active flow control,owing to their distinct advantages,including the absence of moving mechanical parts,a lightweight structure,and rapid response time.24 pairs of alternating current DBD plasma actuators were arranged on a Davis wing.The flow control effectiveness and drag-reduction performance were systematically evaluated through wind-tunnel force measurements,Particle Image Velocimetry(PIV)flow-field measurements,and actual flight tests.Wind tunnel results indicate that plasma actuation effectively modulates near-wall flow structures.The wall jet induced by the actuation in-teracts with near-wall structures,leading to changes in coherent structures:compression of the streamwise extent,enlargement of the spanwise streak spacing,and a reduction in the inclination angle from 15.94° to 9.20°.This flow-control effect suppresses the lift-up motion of quasi-streamwise vortex pairs,weakening momentum transport and at-tenuating Reynolds shear stress in the near-wall region,thereby reducing skin-friction drag.Based on these findings,circling flight tests were conducted using an unmanned aerial vehicle at a fixed altitude and airspeed.Drag reduction effects were evaluated by monitoring changes in motor power under conditions of an airspeed of 24 m/s,a peak-to-peak actuation voltage of 10 kV,and an angle of attack of 4°(close to the maximum lift-to-drag ratio state).The re-sults demonstrate that upon activating plasma actuation,the aircraft's ground speed increased by approximately 7.1%,with a peak drag reduction of 7.59%and an average drag reduction of 6.15%.The consistent drag-reduction trends observed in both wind tunnel and flight tests validate the effectiveness and engineering feasibility of this flow-control method in real-world flight environments.

Over-expansion flow can generate asymmetric shock wave interactions, which lead to significant lateral forces on a nozzle. However, there is still a lack of a suitable theory to explain the phenomenon of asymmetry. The current work carefully investigates the configurations of shock wave interactions in a planar nozzle, and proposes a theoretical method to analyse the asymmetry of over-expansion flows. First, various possible flow patterns of over-expansion flows are discussed, including regular and Mach reflections. Second, the free interaction theory and the minimum entropy production principle are used to analyse the boundary layer flow and main shock wave interactions, establish the relationship between the separation shock strength and separation position, and predict asymmetric configurations. Finally, experiments are conducted to validate the theoretical method, and similar experiments from other studies are discussed to demonstrate the effectiveness of the proposed method. Results demonstrate that the direction of asymmetric over-expansion flow is random, and the separated flow strives to adopt a pattern with minimal total pressure loss. Asymmetric interaction is a mechanism through which the flow can achieve a more efficient thermodynamic balance by minimising entropy production.

An experimental study is conducted on the surface shear stress vector distribution on a plate in a supersonic jet flow, with a focus on the separation region. The shear-sensitive liquid crystal coating (SSLCC) technique is employed for the flow visualization and measurement, which is based on the shear stress distribution, and the flow pattern on the plate is captured. The results demonstrate that the nozzle pressure ratio (NPR) is the main inducement to flow evolution, and a high NPR causes a separation region on the plate, where the adverse flow is challenging to the SSLCC technique. Therefore, an improved measurement method for the SSLCC is proposed to successfully obtain the wall shear stress distribution inside the separation and reattachment area. The flow structures on the plate, including the separation and reattachment positions and vortex and adverse flows, are accurately captured in detail, which indicates that this method is practical for measuring the wall shear stress in separated flow.

Oblique shock train is a common phenomenon in supersonic internal flow, where the mechanism of the asymmetric pattern is still unknown. This paper examines the asymmetric oblique shock train in a Mach number 2.5 flow, and the transition process between top large separation (TLS) and bottom large separation (BLS) has become a key focus of research. The three-dimensional structure of the shock train is investigated from both [Formula: see text] and [Formula: see text] plane perspectives based on schlieren images, which, respectively, exhibit the separation pattern of the shock train on the top/bottom and side walls. Accordingly, the “self-rotation” of the shock train is proposed to explain the TLS-to-BLS transition phenomenon. Experimental results demonstrate that TLS-to-BLS transition is not a two-dimensional process. The TLS-to-BLS transition experiences an intermediate mode, named side large separation (SLS) in this paper, which is similar to a symmetric pattern, while in fact, it is an asymmetric pattern from another plane’s perspective. SLS mode indicates that the three-dimensional shock train holds an asymmetric pattern and rotates 90° around the flow direction; then SLS switches to BLS by rotating another 90°. Consequently, the TLS-to-BLS transition of oblique shock train can be attributed to 180° self-rotation around the flow direction.

Shock Wave/Boundary Layer Interaction is a fundamental problem in hypersonic flows. This study investigates the influence of a moving wall on the flow separation region in a Mach 5 hypersonic flow using high-speed schlieren imaging and high-frequency pressure measurements. The results indicate that the moving wall alters the length and height of the separation bubble and reveal that low-frequency modes dominate the flow evolution between the wedge and the plate, and high-frequency modes capture the influence of the leading-edge shock. The moving wall's influence has three distinct stages. The interaction among the incident shock, the separation shock, and the separation bubble governs the first two stages. The third stage involves the interaction with the leading-edge shock. The differing states of the boundary layer on the moving wall led to distinct trends in the evolution of separation length and height between the first and second stages. Pressure measurements show a consistent trend of pressure fluctuation at various locations on the plate, although the duration and rate of change differ. For a fixed wedge angle, these variations in pressure response are attributed to the different states of the boundary layer on the moving wall. Finally, this paper gives a theoretical model describing the temporal evolution of the separation bubble's length and height based on the free interaction theory and experimental data. These findings offer new insights into controlling shock boundary layer interaction in high-speed flows through a moving wall.

This study investigates the complex multi-scale and nonlinear flow phenomena of reusable rockets during vertical recovery under high subsonic inflow conditions. It focuses on the effects of angle of attack, inflow Mach number, and retro-propulsion jet interference on the aerodynamic characteristics of the rocket airframe. Proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) methods are employed to analyze the unsteady characteristics of the separation zone downstream of the windward end face. The results show that the retro-propulsion jet fundamentally transforms the aerodynamic profile of the “flat-headed cylinder” rocket body. It directly influences the surface pressure distribution on the windward end face and sidewalls, reducing aerodynamic drag and normal force loads. The maximum aerodynamic drag reduction achieved in this study was 67.3 %. At high subsonic conditions, the interaction between the retro-propulsion jet and the inflow induces intense fluctuations in the jet wake region. Pressure disturbances propagate downstream, altering the ambient pressure near the nozzle and causing fluctuations and switching of the shock wave reflection structures at the nozzle exit. Moreover, the retro-propulsion jet significantly modifies the extent and frequency characteristics of the separation zone, transforming vortex shedding and shear layer fluctuations from high-frequency, small-scale patterns to low-frequency, large-scale ones. Overall, this study elucidates the intricate interaction mechanisms between the retro-propulsion jet and subsonic inflow, including flow field reconstruction, pressure distribution changes, and the evolution of separation zone frequency characteristics. These findings offer critical theoretical insights and experimental evidence for the aerodynamic design, performance optimization and flight control of reusable rockets during vertical recovery.

Sliding dielectric barrier discharge (SL-DBD) has demonstrated significant potential for flow control to enhance the aerodynamic performance of aircraft. The paper investigates the velocity characteristics of SL-DBD-induced flow structures during the non-starting phase in continuous mode using particle image velocimetry. The results reveal that the induced flow exhibits periodic velocity fluctuations during the non-starting phase, and the dominance of rotation and strain rates determines the flow structure characteristics. When the rotation rate dominates, the flow structure exhibits vortex behavior. Moreover, the dominance of the rotation rate is not determined by the starting vortex but is an inherent feature of the induced flow itself. When the strain rate dominates, the flow tends to stabilize. However, the rotation rate still plays a role, and the dominance of the rotation and strain rates alternates periodically. By simplifying the flow during the non-starting phase to a dynamic equilibrium state, we derived an expression for the velocity field variation, revealing that the electric field's divergence determines the velocity field's divergence. When the peak-to-peak voltage is 18 kV, the high-frequency fluctuation frequency of ux is 50 Hz, more significant than the high-frequency fluctuation frequency of uy, which is 20 Hz. The investigation demonstrates that the electric field intensity is the primary factor driving the periodic alternation between rotation rate and strain rate, and the fluctuation characteristics of the induced flow velocity are inherent properties. These findings have important implications for understanding the physical mechanisms of SL-DBD-based flow control.

Long-endurance hypersonic vehicles face the dual challenge of withstanding extreme aerodynamic heating while meeting onboard power requirements. Integrating thermoelectric generators within thermal protection systems offers a solution by converting thermal loads into electrical power. However, accurate prediction requires resolving coupled multiphysics, where three-dimensional simulations are computationally prohibitive and existing one-dimensional models lack accuracy. This study develops a quasi-two-dimensional distributed thermal network incorporating shape-factor corrections for rapid, high-fidelity prediction. Multi-objective optimization is performed to balance specific power, thermal expansion mismatch, and thermal margin. Analysis reveals fundamental trade-offs: a maximum-power design achieves 28.1 W/kg but only a 0.8% thermal margin, whereas a balanced design delivers 24.5 W/kg with a 5.1% thermal margin and significantly reduced thermal stress. Despite geometric variations, peak conversion efficiency converges to approximately 13%. This indicates that efficiency is primarily governed by material properties, while geometric optimization effectively tunes temperature and thermal strain distributions, providing guidelines for reliable system development.

An oblique shock train generally forms an asymmetric structure in a Mach-2.7 flow field within a duct. To study the flow structure and interaction between oblique shock trains and upstream shocks, a ramp with equal width was installed inside a Mach-2.7 straight duct to generate an incident shock and an oblique shock train interaction. A Schlieren system, transient pressure measurements and particle image velocimetry were used to capture quantitative and qualitative shock structure information. Results show that the asymmetric separation deflection of the oblique shock train occurs randomly in the symmetrical straight duct. The separation deflection of the oblique shock train was steady with upstream shock interactions. Under backpressure conditions, the rate of movement of the oblique shock train increases rapidly when it passes through the separation regions generated by the ramp, and the deflection direction of the asymmetric separation may switch. Based on the characteristics of the oblique shock train and upstream shock interaction, a flow control method was used to generate asymmetric upstream flow conditions, providing active control of the oblique shock train deflection direction.

The dielectric barrier discharge (DBD) actuator has the advantages of being lightweight, having no moving parts, ease of use, and fast response, and has received widespread attention in flow control applications. Turbulence boundary layer drag reduction is one of many applications of DBD flow control, but the mechanism of DBD actuator turbulence drag reduction needs further investigation. The effect of DBD excitation on the skin-friction drag of a turbulent boundary layer on a flat plate at different flow speeds was investigated experimentally. The change in skin-friction drag was measured using oil film interferometry, and the velocity distribution within the boundary layer was obtained using a particle image velocimetry system. The results showed that under the action of the plasma actuator, the local skin-friction coefficient was measured to decrease by 49%. Through dynamic mode decomposition, plasma actuators can increase the thickness of the low-velocity region in the boundary layer, reduce the intensity of Q2 and Q4 events, and inhibit the development of coherent structures, thereby achieving drag reduction.

The shock wave–boundary layer interaction (SWBLI) phenomenon was investigated experimentally to explore the heat flux distribution characteristics of SWBLIs under different boundary layer flow regimes and the influence/mechanism of different transition positions on peak heat flux of SWBLI. Experiments were conducted in a flow of Mach number 6. The shock generator deflected the flow by 20 °, 22 °, 25 ° and 30 ° resulting in an oblique shock impinging on a flat plate. The rough elements with different heights were arranged to achieve different transition positions. The influence of the relative position of transition and interference zone on the peak heat flux and its physical mechanism are revealed. The results demonstrated that the transitional SWBLI has a higher peak heat flux than the turbulent SWBLI, however the overshoot phenomenon is not reflected in the peak pressure. The increment of heat flux peak in transitional SWBLI is related to the relative positions of transition and interference zones. The increment of heat flux peak of transitional SWBLI compared to turbulent SWBLI can reach 25.5–38.9% and 51.9–65.1% when transition approaches the separation zone and reattachment zone, respectively. This overshoot phenomenon is caused by the streamwise vortex in transitional and so-called laminar SWBLIs, which enhances the energy exchange in the boundary layer and leads to an increase in heat flux. However, in the turbulent SWBLI, the energy mixing in the boundary layer is not significantly enhanced due to the breakup of the large scale streamwise vortex into turbulent small-scale structure.

The present study investigates the discharge and flow characteristics of a sliding discharge (SD) driven by alternating current (AC) and negative direct current (DC) high voltage in continuous operation and burst-mode actuation in quiescent air. The burst frequency f is set at 20, 40, 50, and 100 Hz with a duty cycle τ fixed at 50%. Different actuation cases exhibit similar discharge morphologies and electrical properties. The results indicate that the flow induced by the horizontal body force generated by the SD undergoes the following stages: formation, intensification, accumulation, and stabilization. Based on the effects of the body force, the evolution of the induced flow field can be divided into three stages: the initial stage (starting-vortex stage), the transition stage, and the final stage. In continuous operation, the transition stage is marked by a complex flow structure, while the final stage is distinguished by a deflecting jet. When the burst frequency f ≤ 50 Hz, the duration of the transition stage increases with the burst frequency, and it becomes transient at f = 100 Hz due to the short voltage input time. Phase-averaged particle image velocimetry results indicate that the final stage of the burst-mode actuation can be categorized into three types mostly based on the interaction of the vortices from the AC and DC electrodes. Compared to the continuous operation, the application of the burst-mode actuation in this study has a shorter transition stage duration, resulting in a more rapid realization of flow control.

To provide sufficient lift during takeoff and landing, large aircraft are equipped with complicated high-lift devices. The use of simple flaps coupled with active flow control (AFC) can achieve lift improvement while reducing mechanical structure and weight. The present study focuses on verifying the feasibility and effectiveness of simple flaps combined with sweeping jet flow control. An experimental study on the AFC of flaps, using sweeping jets, was carried out using a NASA SC(2)-0410 supercritical airfoil wind-tunnel model at Re = 2.0 × 105 (with velocity V = 10 m/s). In the experiment, the wing angle of attack (α) ranged from 3 to 18°, and the flap deflection angle (δ) ranged from 0 to 30°; the aerodynamic characteristics and surface pressure characteristics of the wing at typical working conditions were analyzed. Using sweeping jets to control the flow on the flaps, the momentum coefficients (for three actuator groups) of the jet are 0.8%, 3.6%, and 8.2%, respectively, and the maximum lift coefficient was increased by approximately 33%. The influence of the sweeping jet flow rate on the aerodynamic performance of the airfoil is analyzed. There are two main reasons for the lift coefficient increase caused by sweeping jet flow: an extra suction peak near the flap and a suction peak increase near the leading edge area caused by induced flow.

为了研究等离子体激励对湍流边界层的减阻机理,采用交流介质阻挡放电(AC-DBD)的等离子体激励阵列.试验使用热线风速仪和高频粒子图像测速仪(PIV)获取边界层速度分布,平板模型放置在南京航空航天大学0.8m低速直流风洞试验段中(来流速度为11m/s),激励器布置形式为顺来流放置,主要研究不同激励电压下对湍流边界层的减阻效果.研究结果表明,在施加不同电压等离子体激励后得到了不同程度的减阻效果,减阻效果随电压增大出现先增大后减小的趋势,减阻率最大达5.29％.等离子体激励减小了边界层对数区的速度分布,抑制相干结构的发展,削弱相干结构的强度,从而达到减阻的效果.

In this paper, wind tunnel experiments are conducted to study the control law and mechanism of oscillating jet flow control to improve the aerodynamic characteristics of the vertical tail when a civil aircraft encounters left side gust or significant crosswind during takeoff and landing. We measured the vertical tail scaling model’s aerodynamics, spatial flow field, and surface pressure when the Reynolds number was 2.12 × 105. The maximum momentum coefficient of the oscillating jet actuator reaches 0.332%. In addition, we studied the flow control effect of the three-dimensional vertical tail scaled model in different spanwise positions. The experimental results show that the oscillating jet at the rear edge of the stabilizer can significantly increase the lateral force of the vertical tail, and the increment of the lateral force can reach 36.5% under the worst condition of the negative side slip angle of the vertical tail. We can improve the lateral force coefficient of the vertical tail model by applying flow control alone at different spanwise locations. The wing root’s control effect and the vertical tail’s middle section are better than the wing tip’s. The oscillating jet can effectively restrain the flow separation on the rudder. In addition, the input of a high-energy jet “ejects” the mainstream, which increases the flow velocity at the side of the vertical tail actuator. It increases the circulation of the vertical tail. The oscillating jet flow control technology can effectively improve the vertical tail’s steering efficiency and increase the vertical tail’s lateral force, which is of great significance in improving the safety and economy of civil aircraft.

The behavior of a nanosecond pulsed sliding discharge plasma actuator with the ambient pressure from 27 to 101 kPa is experimentally investigated. The electric characteristics, discharge morphologies, and surface temperature distribution of the actuator supplied by constant voltages are studied under different pressure conditions. The threshold pressure for sliding discharge establishment is 54–75 kPa in this paper. The results show that the positive peak value of the current tends to increase as the pressure decreases. Time-integrated discharge images indicate that plasma luminosity and uniformity are enhanced under low-pressure conditions. The discharge morphology of the actuator at 27 kPa displays an alternating distribution of bright and dark stripes. The infrared thermal results demonstrate that decreasing the pressure increases the actuator's surface temperature, intensifies the intensity of spanwise temperature oscillations, and improves the homogeneity in the streamwise direction.

In this paper, shock train motion in a Mach number 2.7 duct is studied experimentally, and large numbers of schlieren images are obtained by a high-speed camera. An image processing method based on Maximum Correlation Detection (MCD) is proposed to detect shock train motion from the schlieren images, based on which the key structures, e.g., separation positions and separation shock angles on the top and bottom walls, can be analysed in detail. The oscillations of the shock train are generated by rhombus and ellipse shafts at various excitation frequencies. According to the analysis of MCD results, the distributions of the frequency components of shock train oscillation generated by the two shafts are distinctly different, in which the motion generated by the ellipse shaft is much smoother; shock train motion is mainly characterized by the oscillation of separation position while the separation shock strength is not so sensitive to downstream disturbance; there is a hysteresis loop relation between the downstream pressure and separation position.

Dielectric barrier discharge (DBD) has been one of the most promising techniques for flow control, but the practical application needs a large plasma surface. A three-electrode sliding discharge plasma actuator (SDPA) can generate plasma filling the inter-electrode distance. In this study, the performance of an SDPA for practical flight at low pressures is reported. When the pressure decreases, plasma discharge becomes more intensive. Current peaks and power consumption of electrode 1 become higher at low pressure. Sliding discharge is fully developed at 54 kPa, and the plasma morphology resembles plumes. PIV results show vortex structures and an induced wall jet above the actuator surface. A ‘potential-arc-discharge’ is observed at 38 kPa, leading to an inhomogeneous surface temperature distribution and a velocity decrease of the jet.