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This paper reports an experimental study on the supersonic jet surface flow structure visualization and shear stress field measurement issuing from a rectangular nozzle with extended shelf. The evolution of the near-field surface flow structures with an increased Nozzle Pressure Ratio (NPR) is successfully captured by the surface oil flow, infrared detection technology, and the Shear-Sensitive Liquid Crystal Coating (SSLCC) technique. Results reveal that under smaller NPR, the wall flow structure is similar to that of a jet without the extended shelf i.e., clean jets, and this is caused by insufficient effect on the boundary layer. However, at higher amplitudes of NPR, there exists a significant effect of the boundary layer, as a near triangular separation forms on the trailing edge of the Mach stem due to the adverse pressure gradient, which is visualized for the very first time in this paper. Furthermore, the vector field of shear stress is measured quantitatively by SSLCC technique. Results shows that the magnitude of shear stress heightened with NPR increasing, and the directions of shear stress changes across the shock wave and expansion fans. In addition, surface streamlines measured by SSLCC is significantly consistent with the streamlines visualized using the oil flow technique.

In this paper a novel design method of aerodynamic configuration is proposed to integrate forebody, strut and inlet for strutjet engine, and a model at design point of Mach number 6 is generated to investigate the aerodynamic performance by both simulations and experiments. The basic flow field employed by proposed method is a combined flow named IBB, which is combined by Internal Conical Flow A (ICFA), truncated Busemann flow I (BI) for external section, and truncated Busemann flow II (BII) for internal section. The model configuration is generated by streamline tracing method from basic flow field, in which the forebody section is traced from ICFA and BI flows, and the inlet as well as strut section is traced from BII flow. The simulations in Mach number 4, 5, and 6 demonstrate uniform starting flow fields with relatively high total pressure recovery, which agree well with experiments in wind tunnel. Additionally, in low Mach number cases, this inlet could start at Mach number 3 while it is unstarted at Mach number 2.7; in high Mach number cases, a uniform flow could still exist in Mach number 6.5 while a relatively strong shock wave boundary layer interaction is found in cowl area of Mach number 7 case, indicating the inlet designed by proposed method works in a relatively wide Mach number range.

The internal drag of the unconventional inner flow channel of blended wing body aircraft must be measured accurately to correct the air intake effect of the blended wing body flow-through model in wind tunnel tests. In this study, the pressure distribution of the inner flow channel under the interaction of internal and external flows was obtained through numerical simulation. A specialized pressure-measuring rake was designed based on the numerical results, and a validation test was conducted in a 2.4 m × 2.4 m transonic wind tunnel. Compared with the flow in traditional inlets/nozzles, the flow in the unconventional inner channel in the current research is asymmetric, the distortion index is higher, and the internal drag is more sensitive to flow changes. The wind tunnel test results have a good correlation with the numerical results, and the repeatability of the test results is satisfactory, indicating that the measurement accuracy and precision of the pressure-measuring rake are acceptable. The design method of the specialized rake is feasible, and it can be used to guide the measurement of complex flow in unconventional inner flow channels of blended wing body aircraft.

In this paper, the incident shock–separation shock interactions on a surface plate near the leading edge are studied theoretically and experimentally, and the transition from regular reflection (RR) to Mach reflection (MR) is the main focus. The theoretical method employs free interaction theory (FIT) and the minimum entropy production (MEP) principle to analyse the separation shock strength of flow separated from the boundary layer and separated from the leading edge, respectively, the criterion based on the MEP principle is employed to predict the RR-to-MR transition near the leading edge. The experiments were performed on a rotatable wedge situated over a sharp leading-edge plate such that the wedge could continuously change the flow deflection angle from by means of a high-precision control device. Fast-response transducers and a high-speed camera were used to measure dynamic pressures and to take schlieren images, respectively. The influences of wedge positions, Reynolds numbers and Mach numbers on shock reflections are investigated by careful tests. The theoretical and experimental results for Mach numbers 5, 6, 7 and 8 show good agreement, indicating that the theoretical method is applicable.

In this work, the Direct Numerical Simulation (DNS) and Oil-Film Interferometry (OFI) technique are used to investigate the hypersonic boundary layer transition induced by single and double roughness elements at Mach number 5. For single roughness, the DNS results showed that both horseshoe vortices and hairpin vortices caused by shear layer instability can affect the boundary layer instability. The generation of the near-wall unstable structure is the key point of boundary layer transition behind the roughness element. At the downstream of the roughness element, the interaction between horseshoe vortices and hairpin vortices will spread in the spanwise direction. For double roughness elements, the effect of the spacing between roughness elements on the transition is studied. It is found that the case of higher spacing between roughness elements is more effective for inducing transition than the lower one. The interaction between two adjacent roughness elements can suppress the evolution of horseshoe vortices in the downstream and trigger the instability of shear layer. Thus, the transition will be suppressed accordingly. Keywords: Boundary layer transition, Hypersonic, Direct numerical simulation (DNS), Oil-film interferometry, Roughness elements

The parallel modular inlet is a common configuration in a strutjet engine of the rocket-based combined-cycle hypersonic vehicle. To study the flow interaction between two adjacent inlets, the flowf...

Prediction of boundary layer transition is important for the design of hypersonic aircrafts. The study of boundary layer transition of hypersonic flow around a flat plate using oil-film interferometry was investigated at Φ500mm traditional hypersonic wind tunnel. In order to measure the skin friction fast and precisely on the hypersonic wind tunnel, the traditional oil-film interferometry technique is improved. A high-speed camera is used to capture the images of fringes and the viscosity of the silicon oil is modified according to the wall temperature measured by thermocouples during the test. The skin frictions of smooth surface and the surface with single square roughness element were measured. For the smooth surface, the boundary layer is laminar. However, the boundary layer transition is promoted by wake vortices induced by the roughness element. Both the results of skin friction with and without the roughness element are in good agreement with the simulation results correspondingly, indicating high accuracy of the oil film interferometry technique.

In this paper the configurations of shock wave-boundary layer interactions (SWBLI) are studied theoretically and experimentally in Mach number 2 and 2.5 flows on test models with various wedge angles ranging from to. The proposed theoretical method couples the free interaction theory (FIT) with the minimum entropy production (MEP) principle to predict the appearance of separation shock, resulting in convex, straight and concave separation shock waves according to different solution combinations, which agree well with current experiments. Additionally, several influences on SWBLI are studied experimentally, in which the parameters related to theoretical solutions are found mostly determining the flow configuration, and SWBLI is much more sensitive to incident shock strength than incoming flow properties. Separation could be suppressed by incident shock when the MEP solution is smaller than the FIT, while it could be intensified when the MEP solution is larger than FIT; by contrast, the effects of separation position and model mounting height could be very weak.

In this paper separation-induced shock reflection is studied theoretically and experimentally. An analytical model is proposed to establish the connections among upstream conditions, downstream conditions and shock configurations. Furthermore, the minimum entropy production principle is employed to determine the incident shock angles as well as the criterion for the transition from regular reflection to Mach reflection, which agrees well with experimental results. Additionally, a solution path for a reflected shock that fulfills the minimum entropy production principle is found in the overall regular reflection domain, based on which the steadiest shock configuration may be determined according to upstream and downstream conditions.

The aim of present investigation is to analyze the unsteady oblique shock train and boundary layer interactions during the self-excited and forced oscillation. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to the forced oscillation. A Schlieren system as well as transient pressure measurements and particle image velocimetry have been used to capture quantitative and qualitative shock structure information. Results show that the behaviors of unsteady SBLIs structure are highly related to the dynamics of shock motion. For both self-excited oscillation and forced oscillation, the asymmetrical characteristics of first X-shock was found to be negatively correlated with shock velocity. There exist some relative motions between the first X-shock and the second shock, but the absolute variations are very weak. At lower excitation frequency, the relative motion is not noticeable to the oscillation amplitude, it could be treated as a rigid motion in the duct. At higher excitation frequency, the relative motion amplitude is significant to the oscillation amplitude, and the relative movement of shock cells becomes the dominant motion. There is a hysteretic effect and phase lag between the shock position and downstream pressure perturbation when the shock train travels along a different path for upstream and downstream movements, and the hysteretic effect becomes weaker with increasing frequency.

An experimental investigation of the response of an oblique shock train to downstream periodic pressure perturbations was conducted. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to a periodic oscillatory motion of the oblique shock train. Six cases of perturbation frequency are studied. The results indicate that the downstream pressure perturbations propagate upstream to cause the oblique shock train to oscillate with a translational motion back and forth and the wall pressure to fluctuate with the same frequency. There is no distinct relative motion between the first oblique shock and the second shock during the motion process of the oblique shock train. The entire oblique shock train exhibits its behaviour of rigid motion and the strength of the first oblique shock of the oblique shock train is nearly stable during its periodic motion. There is a clear hysteresis effect in that the oblique shock train travels along a different path for the upstream and downstream motions. A simple analytical model was built based on these experimental data to analyse the oblique shock train dynamics.

The surface dielectric barrier discharge plasma actuator driven by nanosecond pulses is recognized as an effective fluid actuator for flow separation control. The operation condition of nanosecond dielectric barrier discharge actuators for separated flow control still requires further study, particularly prioritizing the improvement of the effectiveness and reducing energy consumption in flow separation control implementation. In this study, experiments are conducted using a two-dimensional NASA SC(2)-0712 airfoil in a wind tunnel with a Reynolds number of 0.5 × 106 (25 m/s). The pressure measurement experiments show that the location of actuators affects the efficiency of separation control. Particle image velocimetry results indicate that the most efficient location of the actuator is upstream of the separation point and near the original point of the separated shear layer. Meanwhile, the particle image velocimetry results show the vorticity attaches to the airfoil wall after discharge, which suggests that the reattachment is due to the generation of large-scale vortices. These present structures result in the mixing of the shear layer with the main flow thereby delaying separation and reattaching a separated flow. This study shows the most efficient location related to the separation point. Furthermore, it indicates the reattachment of flow is attributed to the motion of vortexes coherent structure.

The modern high performance air vehicles are required to have extreme maneuverability, which includes the ability of controlled maneuvers at high angle of attack. However, the nonlinear and unsteady aerodynamic phenomena, such as flow separation, vortices interaction, and vortices breaking down, will occur during the flight at high angle of attack, which could induce the uncommanded motions for the air vehicles. For the high maneuverable and agile air missile, the nonlinear roll motions would occur at the high angle of attack. The present work is focused on the self-induced nonlinear roll motion for a missile configuration and discusses the influence of the strake wings on the roll motion according to the results from free-to-roll test and PIV measurement using the models assembled with different strake wings at α = 60°. The free-to-roll results show that the model with whole strake wings (baseline), the model assembled with three strake wings (Case A) and the model assembled with two opposite strake wings (Case C) experience the spinning, while the model assembled with two adjacent strake wings (Case B), the model assembled with one strake wing (Case D) and the model with no strake wing (Case E) trim or slightly vibrate at a certain “×” rolling angle, which mean that the rolling stability can be improved by dismantling certain strake wings. The flow field results from PIV measurement show that the leeward asymmetric vortices are induced by the windward strake wings. The vortices would interact the strake wings and induce crossflow on the downstream fins to degrade the rolling stability of the model. This could be the main reason for the self-induced roll motion of the model at α = 60°.

The basic dynamic characteristic of a nanosecond (NS) pulse discharge actuator in still air is investigated, which introduces viscid and inviscid effects to the flow, namely, heated air and pressure waves, respectively. The efficacy of a dielectric barrier discharge plasma actuator driven by a nanosecond pulse for active flow separation control is investigated experimentally on a NASA SC(2)-0712 aerofoil at Re = 0.5 × 10 6 (25 m/s). The pressure distribution on the aerofoil surface is measured, and the result indicates that the separation is well controlled at a high angle of attack. The PIV test result shows that the forcing frequency selectively affects the flow over the aerofoil, and a higher forcing frequency has a better influence on the leading edge shear layer. Schlieren imaging experiments at Re = 0.1 × 10 6 (5 m/s) show that the main disturbance factor for the nanosecond pulse plasma actuator is the heating effect, which changes the local temperature and density and induces several vortex structures. The vortex structures entrain the outer high-momentum fluid into the shear layer, thereby strengthening the mixing of the shear layer with the main flow and delaying separation or even reattaching a separated flow. The research reveals the major disturbance factor introduced by NS discharge applied to flow separation control. Furthermore, the spatiotemporal disturbance process induced by the pulse discharge is investigated.

Plasma-based flow control is one of the most promising techniques for aerodynamic problems, such as delaying the boundary layer transition. The boundary layer’s characteristics induced by AC-DBD plasma actuators and applied by the actuators to delay the boundary layer transition on airfoil at Ma = 0.3 were experimentally investigated. The PIV measurement was used to study the boundary layer’s characteristics induced by the plasma actuators. The measurement plane, which was parallel to the surface of the actuators and 1 mm above the surface, was involved in the test, including the perpendicular plane. The instantaneous results showed that the induced flow field consisted of many small size unsteady vortices which were eliminated by the time average. The subsequent oil-film interferometry skin friction measurement was conducted on a NASA SC(2)-0712 airfoil at Ma = 0.3. The coefficient of skin friction demonstrates that the plasma actuators successfully delay the boundary layer transition and the efficiency is better at higher driven voltage.

Vortex control is a thriving research area, particularly in relation to flying wing or delta wing aircraft. This paper presents the topological structures of vortex flow on a flying wing aircraft controlled by a nanosecond plasma dielectric barrier discharge actuator. Experiments, including oil flow visualization and two-dimensional particle image velocimetry (PIV), were conducted in a wind tunnel with a Reynolds number of 0.5 × 106. Both oil and PIV results show that the vortex can be controlled. Oil topological structures on the aircraft surface coincide with spatial PIV flow structures. Both indicate vortex convergence and enhancement when the plasma discharge is switched on, leading to a reduced region of separated flow.