Kin Hing Lo

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

Static and Wind-on Performance of Polymer-Based Pressure-Sensitive Paints Using Platinum and Ruthenium as the Luminophore

An experimental study has been conducted to investigate the static and wind-on performance of two in-house-developed polymer-based pressure-sensitive paints. Platinum tetrakis (pentafluorophenyl) porphyrin and tris-bathophenanthroline ruthenium II are used as the luminophores of these two polymer-based pressure-sensitive paints. The pressure and temperature sensitivity and the photo-degradation rate of these two pressure-sensitive paints have been investigated. In the wind tunnel test, it was observed that the normalised intensity ratio of both polymer-based pressure-sensitive paints being studied decreases with increasing the number of wind tunnel runs. The exact reason that leads to the occurrence of this phenomenon is unclear, but it is deduced that the luminophore is either removed or deactivated by the incoming flow during a wind tunnel test.

Flow Characteristics along an Active Jets Equipped Contour Bump in a Supersonic Freestream and its Potential to be Applied on Transonic Aircraft for Drag Reduction: An Experimental Study

Flow separation control stragies on contour bump models under a Mach 1.3 freestream has been investigated experimentally in a transonic/supersonic wind tunnel. Surface oil flow visualisaiton, Schlieren photography, and Particle Image Velocimetries (PIV) measurements have been conducted. Flow separation, in both streamwise and spanwise directions, was observed at the valley of the plain contour bump. Usage of passive by-pass jets on contour bump show no effects in reducing the extent of flow separation in both streamwise and spanwise directions. It is observed that both spanwise and streamwise flow separation are reduced by blowing jets from the valley of the contour bump. It is believed that the jets hinder the formation of the spanwise, counter-rotating vortrices and deflect the shear layer to make the flow at the valley becomes more attach to the surface of the bump. Two active jet pressures of 1 bar and 4 bar have been employed in this study. It is observed that a higher jet pressure is more effective in achieving flow separation control in both streamwise and spanwise directions. Further investigations will be conducted to obtain surface pressure profiles on the contour bump models.

Vehicle and mission design of a future small payload launcher

This paper presents the conceptual design and performance analysis of a partially reusable space launch vehicle for small payloads. The system uses a multi-stage vehicle with rocket engines, with a reusable first stage capable of glided or powered flight, and expendable upper stage(s) to inject a 500 kg payload in different low Earth orbits. The space access vehicle is designed to be air-launched from a modified aircraft carrier. The aim of the system design is to develop a commercially viable launch system for near-term operation, thus emphasis is placed on the efficient use of high TRL technologies. The vehicle design are analysed using a multi-disciplinary design optimisation approach to evaluate the performance, operational capabilities and design trade-offs.

Flow Physics of a Three-Dimensional Rounded Contour Bump in a Mach 1.3 Supersonic Free-stream

An experimental study has been conducted to investigate the flow physics along a three-dimensional contour bump in a Mach 1.3 supersonic free-stream. Monochrome Schlieren photography technique was employed for flow diagnostics. Results obtained from the experimental study indicated that the flow pattern appears over the bump upper surface depends on the shock impingement location. It was observed that two lambda shock structures were formed over the bump upper surface in both cases being studied. It was observed that when the normal shock is impinged at the bump crest, the flow downstream of the bump crest remains relatively attached while separated flow appeared downstream of the bump crest for all other normal shock impingement locations.

Incident Shock-Transverse Jet Interactions at Mach 1.9: Effect of Shock Impingement Location

The scramjet engine is an efficient design for high-speed propulsion, requiring injection of fuel into a supersonic flow in a short amount of time. Due to the nature of the flow numerous shock waves exist within the combustor of a scramjet, significantly altering the flow characteristics and performance of the engine as the flow Mach number or attitude is changed. According to Mai et al. [1] the location of impingement of the incident shock, relative to the fuel injection location, has significant impact on the mixing and flame-holding properties. This emphasises the importance of understanding and hence the need for controlling the dynamic interactions that are created. Of course another fertile area where transverse jet injections are studied for their application is the creation of forces and moments for pitch and attitude control [2, 3].

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.

Flow visualisation of a normal shock impinging over a rounded contour bump in a Mach 1.3 free-stream

An experimental study has been conducted to visualise the instantaneous streamwise and spanwise flow patterns of a normal shock wave impinging over a rounded contour bump in a Mach 1.3 free-stream. A quartz-made transparent shock generator was used, so that instantaneous images could be captured during the oil-flow visualisation experiments. Fluorescent oil with three different colours was used in the surface oil-flow visualisation experiment to enhance the visualisation of flow mixing and complicated flow features that present in the flow field. Experimental data showed that the rounded contour bump could split the impinging normal shock wave into a or a series of lambda-shaped shock wave structure(s). In addition, it was found that the flow pattern and the shock wave structures that appeared over the rounded contour bump depended highly on the impinging location of the normal shock wave. The flow pattern shown in this study agreed with the findings documented in literature. Moreover, it was observed from the instantaneous oil streaks that the normal shock impinging location also affected the size and the formation location of the spanwise counter-rotating vortices downstream of the bump crest. Finally, it was concluded that the terminating shock could distort the oil streaks that left over the surface of the contour bump. Therefore, the use of the transparent normal shock wave generator is recommended when conducting experiments with normal shock wave impingement involved.Graphical abstract

Effectiveness of Flow Separation Control on Contour Bumps under a Mach 1.3 Freestream: An Experimental Study

Flow separation control stragies on contour bump models under a Mach 1.3 freestream has been investigated experimentally in a transonic/supersonic wind tunnel. Surface oil flow visualisaiton, Schlieren photography, and Particle Image Velocimetries (PIV) measurements have been conducted. Flow separation, in both streamwise and spanwise directions, was observed at the valley of the plain contour bump. Usage of passive by-pass jets on contour bump show no effects in reducing the extent of flow separation in both streamwise and spanwise directions. It is observed that both spanwise and streamwise flow separation are reduced by blowing jets from the valley of the contour bump. It is believed that the jets hinder the formation of the spanwise, counter-rotating vortrices and deflect the shear layer to make the flow at the valley becomes more attach to the surface of the bump. Two active jet pressures of 1 bar and 4 bar have been employed in this study. It is observed that a higher jet pressure is more effective in achieving flow separation control in both streamwise and spanwise directions. Further investigations will be conducted to obtain surface pressure profiles on the contour bump models.

Micro-Ramps Flow Characteristics at Mach 1.9 & 5

The prominent problem faced by the air-breathing propulsion system of supersonic aero-vehicles is the phenomenon known as shock wave- boundary layer interactions (SBLI) which cause boundary layer separation and adverse pressure gradients