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

Supersonic flow over rounded contour bumps with vortex generators or passive longitudinal jets

An experimental study has been conducted to investigate the flow characteristics over two rounded contour bumps. Vane-type vortex generators or longitudinal aligned passive by-pass jets were implemented in attempt to achieve wake flow control in rounded contour bumps. According to the results collected from the surface oil flow visualisation experiments, it was observed that the use of both the vane-type vortex generators and the longitudinally aligned passive by-pass jet could reduce the size of the spanwise vortices in the bump valley. In addition, a pair of streamwise horseshoe vortices was observed downstream of the bump crest of the contour bump that equipped with the vane-type vortex generators. From the data collected in the particle image velocimetry measurements, it was found that the use of both the vane-type vortex generators and the longitudinally aligned passive by-pass jet could not reduce the size of the wake region but they could reduce its strength. It is deduced that the two streamwise horseshoe vortices generated by the vane-type vortex generators enhance flow mixing which results in reducing the strength of the wake region. In contrast, blowing passive by-pass jet in the bump valley increases the local flow velocity in order to reduce the strength of the wake region

Flow characteristics of various three-dimensional rounded contour bumps in a Mach 1.3 freestream

Streamwise and spanwise flow pattern over three rounded contour bumps with different flow control strategies employed have been experimentally investigated in a Mach 1.3 freestream. Surface oil flow visualisation, Schlieren photography and particle image velocimetry measurements were used for flow diagnostics. Experimental data showed that in a Mach 1.3 freestream over the baseline plain bump, significant flow separation appeared at the bump crest that led to the formation of a large wake region downstream. In addition, two large counter-rotating spanwise vortices were formed in the bump valley. It was observed that the use of the passive by-pass blowing jet in the bump valley showed no obvious effects in reducing the sizes of both the wake region and the spanwise vortices in the bump valley. In contrast, it was found that the size of the wake region and the spanwise vortices could be reduced by blowing sonic jet in the bump valley. This approach of flow control found to be the most effective when the total pressure of the blowing jet was 2 bar. It is deduced that the active blowing jet hindered the formation of the spanwise vortices in the bump valley as well as deflected the shear layer downwards so that a smaller re-circulating bubble was formed downstream of the bump crest

Shock-Induced Separation of Transitional Hypersonic Boundary Layers

This thesis presents a joint experimental/CFD investigation of shock-induced boundary layer separations in hypersonic transitional boundary layers with an emphasis on collapse and re-establishment times of the separation bubble. This study also provides high fidelity measurements and excellent characterisation of the flow field in order to provide benchmark data of a challenging flow configuration with which to benchmark next generation CFD solvers. The experiments were conducted in the Imperial College Aeronautics Department Number Two Gun Tunnel, a Mach 8.9 axisymmetric facility with a freestream unit Reynolds number of 47 million An axisymmetric blunt-nosed cylinder fitted with an 8 degree flare forms the primary vehicle for this study, although a 1.3 degree cowl geometry was also used to impinge a shock onto the blunt-nosed cylinder.. The shock boundary layer interaction was designed such that it was separated for a laminar boundary layer and collapsed for a turbulent one. Carefully controlled turbulent spots were generated upstream of the interaction region which passed through the separation causing its collapse and subsequent re-establishment. Two intermittency cases are considered, one where turbulent spot spacing is large and collapse/re-establishment pairs can be considered independent of each other and one where they can not. Experimental surface quantities through the interaction region are measured using either heat-transfer or pressure measurements and schlieren video is used to diagnose the larger shock structure. Further a non-intrusive toluene PLIF method is assessed for use in this facility and shows promise. CFD simulations are done using an in-house operator split Godunov solver with a Baldwin-Lomax turbulence model. CFD simulations show good agreement with experiment and provides information on flow quantities that would be extremely difficult to measure otherwise. Collapse times of the separation bubble were found to be fast in relation to characteristic spot passage times. The collapse process is also fast in relation to the surrounding flows ability to adjust, with collapse associated with significant shock curvature of the immediate outboard shock structures. This leads to unsteadiness, with surface pressure measurements exceeding the range bounded by the laminar separated and turbulent collapsed cases. The severity of the unsteadiness appears to be driven by turbulent spot spacing. Re-establishment is considerably slower, showing asymptotic recovery that is likely driven by viscous diffusion rates, taking many characteristic spot passage times to recover.Open Acces

Fluid-structure-interaction of adaptive shock control bumps

Improving aircraft aerodynamic efficiency is key to achieve the desired performance enhancements, in the light of ambitious environmental targets. Flow control measures are required to mitigate the detrimental effects of normal shock waves, which occur on the wings of modern commercial aircraft and generate wave drag. A review of the relevant literature has identified the potential of Shock Control Bumps (SCBs) for reducing overall drag, although only for a narrow operating envelope. Recent work has shown that, through geometric modifications, adaptive SCBs could be effective over a wider range of flow conditions. However, further research is required to fully assess their aero-structural behaviour. In the present study, an investigation of the Fluid-Structure-Interaction of adaptive SCBs has been conducted to characterise their unsteady behaviour and assess their potential for reducing drag. The adaptive SCB was modelled as a thin aluminium alloy flexible plate, deployed by means of different actuation mechanisms, and was tested at Mach 1.4 in the Imperial College supersonic wind tunnel for a range of flow conditions. Point-tracking photogrammetry and Pressure Sensitive Paint (PSP) were successfully applied to the challenging setup of a deforming flexible surface in transonic flow, to obtain reliable full-field deformation and surface pressure measurements during experiments. The accuracy of these techniques and their sensitivity to testing conditions have been estimated. RANS CFD simulations of the flow above a rigid SCB, performed in OpenFOAM, complement experiments. Computed stagnation pressure profiles downstream of the rigid SCB aided the development of a methodology to evaluate the drag reduction potential of adaptive SCBs from experiments in a blow-down supersonic wind tunnel. In addition, stability of the shock travelling in the wind tunnel working section was shown to be related to the drag generated by the bump for various flow conditions. The performance of two- and three-dimensional adaptive SCBs has been assessed and compared with the baseline case of a solid flat plate. Carefully designed adaptive SCBs were found to successfully bifurcate the strong normal shock that is expected on next generation transonic wings, and, through passive geometric modifications, maintain good properties over a range of flow conditions. This study has showed the promising potential of adaptive SCBs for mitigating the off-design performance penalty typical of rigid SCBs, without compromising the on-design drag saving capabilities. Some recommendations for further work are suggested.Open Acces

Shock wave-boundary-layer interactions in high-speed intakes

Shock wave boundary layer interaction (SWBLI) occurs in many aerospace applications such as wings in high-speed flight, missiles, and supersonic intakes. Key to the design of the latter is compressing a large volume of air with SWBLIs while maintaining maximum total pressure recovery and minimum flow distortion over a wide operating range. Under specific conditions, the formation of multiple SWBLIs (shock trains) within the intake can occur. Over the years, many numerical methods and models have been employed to predict the flow physics of shock trains. This work aims to determine the suitability of non-linear RANS turbulence closures for modelling shock trains in ducted geometries by implementing several non-linear closures in the University of Glasgow HMB3 CFD solver. First, the best modelling techniques for matching the experimental conditions were identified by performing validations against several shock train experiments. As a next step, several non-linear RANS closures were implemented in the solver. All closures improved the predicted wall pressures by accounting for the secondary flows present near the duct corners. The closures accounted for the secondary flow by predicting a fair level of normal Reynolds stress anisotropy near the corner of the duct. It was found that even simple non-linear closures based on quadratic constitutive relations result in significant improvements compared to linear closures. Additional simulations were performed at different Mach numbers, Reynolds numbers, and back (exit) pressures to assess the robustness of the non-linear closures and the sensitivity of the solution to changes in modelling parameters. It was observed that the flow distortion decreases rapidly downstream of the first shock in the shock train and that it is greatly influenced by its structure. As a final step, simulations of a shock train in a geometry representative of a highspeed intake were performed to assess the suitability of the closures for practical (real-world) applications. Three different geometries resulted in considerably different shock train structures compared to the ones in ducts. The flow distortion downstream of the shock train was found to be sensitive to both the incidence and roll angles. The SWBLI exhibited upstream and downstream movements within the intake with increasing incidence and roll angles

Directed energy deposition flow control for high speed intake applications

One of the key enabling technologies for high speed airframes is the propulsion system. Some of the important factors in any widely used propulsion system are reliability and efficiency. Both of these performance indicators are areas that can still be improved for a popular high speed propulsion system, the Ramjet/Scramjet. Regarding the propulsion system efficiency, one component that has a significant impact is the intake. Specifically the intake efficiency, often quantified by the total pressure loss, can significantly influence the overall system performance. Now considering the reliability of the system, a phenomenon known as unstart can present issues for a Ramjet/Scramjet during its operation. This phenomenon refers to the ejection of the shock structure, known as a shock train, from within the internal duct in the ramjet/Scramjet called the isolator. This results in a bow shock upstream of the propulsion system intake leading to significant drag increases and thrust decreases due to reductions in intake mass flow. This work aims to propose methods that can tackle both of the issues highlighted above. The goal is to use energy deposition based flow control methods to improve the system performance. To achieve this, the first step was the improvement of the available wind tunnel facilities. Through the careful examination and numerical simulations of the existing wind tunnel design, key issues are identified and solved methodically. This included the re-design of the nozzle; both the subsonic and supersonic sections of the nozzle are carefully designed. The diffuser is also re-designed to ensure it is suitable for the desired test section Mach number. Finally, the location of the wind tunnel relative to the vacuum tank is examined and changes are made in order to reduce the conductance present within the system to a level at which the desired mass flow rate could be achieved, and the wind tunnel can operate at the desired flow conditions. It has previously been shown that the inclusion of a cavity within the isolator can improve the unstart margin. However the introduction of a cavity will result in an increase in the drag of the overall propulsion system. Therefore the aim of this work, with regard to improving the unstart performance of the propulsion system, is to examine how energy deposition flow control can be used to control a supersonic cavity flow. This research will examine the use of nanosecond, high voltage dielectric barrier discharge (ns-DBD) plasma actuators to control the cavity flow. However before this, the interaction of the cavity with the baseline internal duct flow was examined. This was in order to classify the scale of cavity that could be included in a duct before the influence on the flow was too large for the duct to start. Through Schlieren imaging, it is shown that increasing cavity scale led to an increase in the cavity shear layer which, in turn, became a limiting factor in the ability of the supersonic duct flow to start successfully. It is suggested that increasing cavity length is the dimension with the largest impact on the starting of the duct flow. This work allowed the cavity being examined for flow control applications to be appropriately scaled so as to still allow the supersonic duct flow to start correctly. The second topic that is necessary before the cavity flow control can be examined is the charaterisation of the ns-DBD actuators. The key goal of this section of research is to identify the combination of actuators parameters that resulted in the strongest flow control influence, measured in this case by the generated pressure wave strength. This characterisation process is conducted through the image processing of the captured Schlieren images during the actuator operation. It is found that the only parameters investigated with significant impact on the actuator control authority are electrode length and dielectric thickness. Shorter electrodes and thinner dielectric barriers are found to be the most successful, for the same electrical input signal used and the same dielectric material. Whilst characterising the actuators for control authority, the electrical efficiency of the actuators is also investigated using electrical measurements, quantitative Schlieren and infrared thermography. The quantitative Schlieren and electrical measurements were found to give accurate results, however they did not present any clear conclusions regarding how the geometric parameters investigated influenced the electrical efficiency. The infrared thermography was unsuccessful due to the small temperature variations being measured and the relatively large uncertainty of the equipment used. Following the characterisation of the ns-DBD plasma actuators, they were applied as flow control to the supersonic cavity flow. Laser vibrometry, Schlieren and oil flow are used to investigate the impact that these actuators have on the baseline supersonic cavity flow. It is found that the actuators available are not able to provide any measureable flow control authority on the baseline cavity flow. The other performance indicator highlighted above is the efficiency of the intake. This study focuses on the improvement of intake total pressure recovery through the use of laser energy deposition flow control. The aim of the flow control is to mitigate the separation observed on the external compression ramp and, as a result, mitigate the negative impact of the buzz instability, commonly observed in supersonic intakes. It is shown that the use of laser energy deposition can provide an improvement in total pressure recovery and does reduce the impact of these buzz instability modes. It is also illustrated how, for a particular set of example conditions, this could improve the overall fuel consumption of a given Ramjet propulsion system

Supersonic Corner Flows in Rectangular Channels

Rectangular channel geometries are widely encountered in supersonic flows, such as in wind tunnels and in aircraft inlets. Shock-boundary-layer interactions in these flows are known to exhibit significant three-dimensionality, due to the presence of sidewalls and associated corner boundary layers. The main effect is on the local separation of these corner regions, which then affects the wider flow field. Successful prediction of the overall flow therefore relies on the corner separation to be determined accurately. This, in turn, requires knowledge of the flow momentum distribution within the corner boundary layers. However, numerical methods struggle to reliably compute these flows and there is not much experimental data on supersonic corner boundary layers for comparison. This thesis addresses the outstanding gap in knowledge by performing validation-quality experiments on the corner regions of a Mach 2.5 channel flow, with a unit Reynolds number of approximately 40 million per metre. The experiments are conducted in the rectangular test section of a supersonic wind tunnel at the University of Cambridge. An analysis of the wind tunnel experiments, alongside computational data provided by the US Air Force Research Laboratory, reveals that the corner boundary layers are strongly influenced by the geometry of the two-dimensional nozzles used to produce the supersonic flow. The dominant effect is related to bulk vertical velocities within the sidewall boundary layers, induced by vertical pressure gradients in the nozzle. For some very particular geometries, a second influence may be associated with a region of separated flow immediately ahead of the nozzle, which generates vortices within the sidewall boundary layer. Through these mechanisms, the nozzle geometry is seen to strongly influence both the thickness and the structure of the corner boundary layers. High-quality experimental data in the corner regions are used to validate relevant numerical methods. Simple linear eddy-viscosity type turbulence models are found to compute these flows particularly poorly, with a 7% discrepancy in streamwise velocity. This is largely due to the fact that they do not capture known, stress-induced, corner vortices. However, the quadratic constitutive relation improves prediction of the corner boundary-layer structure, reducing experimental-computational differences by as much as half. This improvement is associated with vorticity generation in these corner regions, albeit with slightly different properties to the physical vortices. This production of vorticity depends only on the presence of a quadratic term in the eddy-viscosity model and not on which particular quadratic term is used. A more general form of the quadratic constitutive relation with one additional term is proposed, which appears to exhibit substantial improvements in the prediction of turbulent stress anisotropies. The nozzle geometry effects are exploited to produce two otherwise-identical experimental setups with distinctly different momentum distributions in the corner boundary layers. A full-span wedge introduces an oblique shock with flow deflection angle, 8 degrees, which impinges on the floor boundary layer. The two setups exhibit quite dissimilar separation behaviour, not only in the corner regions but also on the tunnel's centre span, with a difference in central separation length of as much as 35%. The observed behaviour is consistent with expectations based on local flow momentum affecting corner separation size, and on the displacement effect of this corner separation influencing the wider flow.This material is based upon work supported by the US Air Force Office of Scientific Research under award FA9550–16–1–0430