K. Kontis

Control of Flow Separation on a Contour Bump by Jets: An Experimental Study

The usage of three-dimensional contour bumps in shock control under transonic flows becomes an active research topic in the aerospace industry

Experimental Studies on Shock Wave Propagating Through Junction with Grooves

Using a shock tube to generate shock waves, the wave patterns through several bifurcated junctions were studied. The effect of roughness was simulated by introducing grooves on the lower wall of the junctions, while the upper wall had a smooth surface. Five different bifurcation angles were used in the present study: 0, 30, 45, 60, and 90 degrees. The shock tube driver pressure was P4 = 12 bar, and the pressure within the shock tube driven section, P1, was atmospheric. Air was used as the driver and driven gas. High-speed schlieren photography, using the Shimadzu Hyper-Vision camera, was used to visualise the flow-field generated by the propagating incident shock wave. The interval between each frame was of 4 µs. Pressure measurements were taken to quantify the attenuation of the incident shock wave at the exit of the different junctions.

Micro-ramps in Mach 5 hypersonic flow

Shock/boundary layer interaction (SBLI) is an undesirable phenomenon, occurring in high-speed propulsion systems. The conventional method to manipulate and control SBLI is using a bleed system that involves the removal of a certain amount of mass of the inlet flow to control boundary layer separation. However, the system requires a larger nacelle to compensate the mass loss, larger nacelles contribute to additional weight and drag and reduce the overall performance. This study investigates a novel type of flow control device called micro-ramps, a part of the micro vortex generators (VGs) family that intends to replace the bleed technique. Micro-ramps produce pairs of counter-rotating streamwise vortices, which help to suppress SBLI and reduce the chances of flow separation. Experiments were done at Mach 5 with two micro-ramp models of different sizes. Schlieren photography, surface flow visualization and infrared thermography were used in this investigation. The results revealed the detailed flow characteristics of the micro-ramp, such as the primary and secondary vortices. This helps us to understand the overall flow physics of micro-ramps in hypersonic flow and their application for SBLI control.

Experimental Studies on Micro-Ramps at Mach 5

The performance of hypersonic propulsion can be critically affected by shock wave/boundary layer interactions (SBLIs), whose severe adverse pressure gradients can cause boundary layer separation. This phenomenon is very undesirable in engine intakes leading to total pressure loss and flow distortion which can cause engine unstart. Hence, it is essential to apply flow control method to the flow, either at the beginning or during the interaction phenomenon to prevent the shock-induced separation [1].

A note on the generation of a compressible vortex rings using helium as driver gas

An experimental study has been conducted on the generation and propagation of compressible vortex rings using helium as a driver gas, with the aim of evaluating the effects of multi-gas operations for real-life applications. The advantage of such system, when compared to a constant gas system based on ambient air, is to effectively increase the Mach number while keeping the pressure ratio constant. Three pressure ratios of ∼4, 8 and 12 were set, corresponding to experimental Mach numbers of approximately 1.50, 1.81 and 2.05. The increase in incident Mach number resulted in the variation of the vortex ring and trailing jet structure, and an increase in both the velocity magnitude and vorticity field. Results showed a transition from the regular-reflection shock-cell system at the experimental Mach number approximately 1.50 to the presence of a Mach reflection with a central Mach disc, which grew in size with further increase in incident Mach number. The presence of the Mach disc resulted in the formation of a subsonic jet, internal to the main trailing jet. Its velocity was measured to be in the order of magnitude of 550 m/s, with the speed of sound of helium at 1005 m/s. Results also demonstrated that shear layers formed between the subsonic and main trailing jet have opposing vorticity, with that of the subsonic jet being approximately half in magnitude. Secondary counter-rotating vortex rings were generated ahead of the main vortex and orbited around it. The analysis of the vorticity field has shown that these secondary vortices have a magnitude approximately half of that of the vorticity of the main vortex, and has confirmed that they have an opposite direction of rotation.

Experimental study of dual injections with a cavity in supersonic flow

The design of supersonic injection systems is a key issue for the development of a supersonic combustion ramjet (scramjet) [1]. Since supersonic flow is present in the combustion chamber of the scramjet, it is difficult to satisfy features which indicate fuel-air mixing and flame stabilisation.

Application of Pressure- and Temperature-Sensitive Paint in a Hypersonic Double Ramp Flow

Pressure- and Temperature-Sensitive Paint (PSP, TSP) was considered as novel nonintrusive flow diagnostics which can provide quantitative global surface pressure and temperature mapping on complex geometry with high spatial resolution [1]. PSP and TSP has been mainly applied in transonic flows [2, 3, 4], supersonic flows [5, 6, 7] and low speed flows [8]. The application of PSP and TSP has been extended to hypersonic flow [9, 10, 11] but not reported as many as at the other flow conditions. The testing condition is challenging for PSP measurement in hypersonic mainly based on the following reasons. The critical aerodynamic heating will cause considerable error for the PSP measurement because of thermal quenching mechanism.

Effect of Roughness in Jets in Mach 5 Cross Flow

Transverse jet injection into supersonic/hypersonic cross flow has been encountered in many engineering applications ranging from scramjet combustors and solid rocket motor or liquid engine thrust vector control systems to high speed flying vehicle reaction control jets. These applications all involve complex three dimensional flow patterns comprising separated regions, shock waves, shear layers and wakes in common. Owing to numerous applications and these complicated flow features, transverse injections over different geometries and various forebodies have been received significant amount of interest. Earlier studies were focused on wind tunnel experiments and the utilisation of conventional measurement techniques such as Schlieren/Shadowgraph photography, wall pressure and concentration measurements to better understand the jet interaction and penetration phenomena. These studies aimed to assess the effect of injection pressure ratio, location of injection and state of incoming boundary layer and type of injectant gas on jets in supersonic/hypersonic cross flow. Recent studies of missile/forebody applications involving reaction control jets by several researchers have investigated jet interaction phenomenon on various axisymmetric body configurations at supersonic/ hypersonic speeds [1, 2, 3, 4]. Their aim was to investigate control effectiveness of transverse/lateral jets on different missile body configurations.

Interaction between Laser Induced Plasma and Boundary Layer over a Flat Plate in Hypersonic Flow

Laser energy deposition has brought great interest to researchers due to its applications in drag reduction [1, 2], shock wave modification [3, 4], fuel ignition [5], and optical perturber for transition study [6, 7, 8]. Compared to the electric discharge flow control techniques, laser energy deposition can excite the flow non-intrusively with almost any pulse-width and repetition rate [9] with out any electrodes.