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The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved. In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles. Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed. A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch. The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation. Keywords:turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.

Boundary layer separation is an unwanted phenomenon in most technical applications, as for instance on airplane wings, ground vehicles and in internal flow systems. If separation occurs, it causes loss of lift, higher drag and energy losses. It is thus essential to develop methods to eliminate or delay separation.In the present experimental work streamwise vortices are introduced in turbulent boundary layers to transport higher momentum fluid towards the wall. This enables the boundary layer to stay attached at  larger pressure gradients. First the adverse pressure gradient (APG) separation bubbles that are to be eliminated are studied. It is shown that, independent of pressure gradient, the mean velocity defect profiles are self-similar when the scaling proposed by Zagarola and Smits is applied to the data. Then vortex pairs and arrays of vortices of different initial strength are studied in zero pressure gradient (ZPG). Vane-type vortex generators (VGs) are used to generate counter-rotating vortex pairs, and it is shown that the vortex core trajectories scale with the VG height h and the spanwise spacing of the blades. Also the streamwise evolution of the turbulent quantities scale with h. As the vortices are convected downstream they seem to move towards a equidistant state, where the distance from the vortex centres to the wall is half the spanwise distance between two vortices. Yawing the VGs up to 20° do not change the generated circulation of a VG pair. After the ZPG measurements, the VGs where applied in the APG mentioned above. It is shown that that the circulation needed to eliminate separation is nearly independent of the pressure gradient and that the streamwise position of the VG array relative to the separated region is not critical to the control effect. In a similar APG jet vortex generators (VGJs) are shown to as effective as the passive VGs. The ratio VR of jet velocity and test section inlet velocity is varied and a control effectiveness optimum is found for VR=5. At 40° yaw the VGJs have only lost approximately 20% of the control effect. For pulsed VGJs the pulsing frequency, the duty cycle and VR were varied. It was shown that to achieve maximum control effect the injected mass flow rate should be as large as possible, within an optimal range of jet VRs. For a given injected mass flow rate, the important parameter was shown to be the injection time t1. A non-dimensional injection time is defined as t1+ = t1Ujet/d, where d is the jet orifice diameter. Here, the optimal  t1+ was 100-200. / QC 20100825

The separating turbulent flow in a plane asymmetric diffuser with 8.5 degrees opening angle is investigated experimentally and computationally. The considered flow case is suitable for fundamental studies of separation, separation control and turbulence modelling. The flow case has been studied in a specially designed wind-tunnel under well controlled conditions. The average velocity and fluctuation fields have been mapped out with stereoscopic particle image velocimetry (PIV). Knowledge of all velocity components allows the study of several quantities of interest in turbulence modelling such as the turbulence kinetic energy, the turbulence anisotropy tensor and the turbulence production rate tensor. Pressures are measured through the diffuser. The measured data will form a reference database which can be used for evaluation of turbulence models and other computational investigations. Time-resolved stereoscopic PIV is used in an investigation of turbulence structures in the flow and their temporal evolution. A comparative study is made where the measured turbulence data are used to evaluate an explicit algebraic Reynolds stress turbulence model (EARSM). A discussion regarding the underlying reasons for the discrepancies found between the experimental and the model results is made. A model for investigations of separation suppression by means of vortex generating devices is presented together with results from the model in the plane asymmetric diffuser geometry. A short article on the importance of negative production-rates of turbulent kinetic energy for the reverse flow region in separated flows is presented. A detailed description of the experimental setup and PIV measurement procedures is given in a technical report. / QC 20100923

<p>The increasing demand for higher performance in rocketlaunchers promotes the development of nozzles with higherperformance, which basically is achieved by increasing theexpansion ratio. However, this may lead to flow separation andensuing instationary, asymmetric forces, so-called side-loads,which may present life-limiting constraints on both the nozzleitself and other engine components. Substantial gains can bemade in the engine performance if this problem can be overcome,and hence different methods of separation control have beensuggested. However, none has so far been implemented in fullscale, due to the uncertainties involved in modeling andpredicting the flow phenomena involved.</p><p>In the present work the causes of unsteady and unsymmetricalflow separation and resulting side-loads in rocket enginenozzles are investigated. This involves the use of acombination of analytical, numerical and experimental methods,which all are presented in the thesis. A main part of the workis based on sub-scale testing of model nozzles operated withair. Hence, aspects on how to design sub-scale models that areable to capture the relevant physics of full-scale rocketengine nozzles are highlighted. Scaling laws like thosepresented in here are indispensable for extracting side-loadcorrelations from sub-scale tests and applying them tofull-scale nozzles.</p><p>Three main types of side-load mechanisms have been observedin the test campaigns, due to: (i) intermittent and randompressure fluctuations, (ii) transition in separation patternand (iii) aeroelastic coupling. All these three types aredescribed and exemplified by test results together withanalysis. A comprehensive, up-to-date review of supersonic flowseparation and side-loads in internal nozzle flows is givenwith an in-depth discussion of different approaches forpredicting the phenomena. This includes methods for predictingshock-induced separation, models for predicting side-loadlevels and aeroelastic coupling effects. Examples are presentedto illustrate the status of various methods, and theiradvantages and shortcomings are discussed.</p><p>A major part of the thesis focus on the fundamentalshock-wave turbulent boundary layer interaction (SWTBLI) and aphysical description of the phenomenon is given. Thisdescription is based on theoretical concepts, computationalresults and experimental observation, where, however, emphasisis placed on the rocket-engineering perspective. This workconnects the industrial development of rocket engine nozzles tothe fundamental research of the SWTBLI phenomenon and shows howthese research results can be utilized in real applications.The thesis is concluded with remarks on active and passive flowcontrol in rocket nozzles and directions of futureresearch.</p><p>The present work was performed at VAC's Space PropulsionDivision within the framework of European spacecooperation.</p><p><b>Keywords:</b>turbulent, boundary layer, shock wave,interaction, overexpanded,rocket nozzle, flow separation,control, side-load, experiments, models, review.</p>

This thesis presents the experimental results on the formation of flow separation from a rigid wall with a salient edge. In the case of automotive vehicles or aircrafts with rear cargo compartment doors, such salient edges are at the origin of separated wake flows resulting in increased drag and other disturbing effects. Recent studies of Ahmed et al. (1984) on simplified geometries showed the strong influence of the slant angle on the flow separations. In this study, the geometry is further simplified to examine the flow separation under two-dimensional conditions. The experimental configuration consists of a fixed horizontal front panel and an attached rear panel with variable slant angle. The experiments were carried out in a low speed water channel to analyze the flow structure by flow visualization techniques. The hydrogen bubble technique nd PIV measurements are used to obtain both qualitative and quantitative information on the flow structure. The electrolytic precipitation technique is used to analyze the flow separation in more detail. The slant angle varied between 0 and 35 degrees while the Reynolds numbers of the model was fixed to 24800 and 50500. As a function of slant angle and Reynolds number, two different types of flow separation were observed: boundary layer separation due to adverse pressure gradient and the so called &ldquo / inertial separation&rdquo / at the edge singularity. Future strategies to control the formation of the wake flow highly depend on the very different flow structure of these two types of separation.

The objective of this thesis is to accomplish the three dimensional parallel thin-layer Navier-Stokes solutions for low aspect ratio rectangular flat wings in compressible flow. Two block parallel Navier Stokes solutions of an aspect ratio 1.0 flat plate with sharp edges are obtained at different Mach numbers and angles of attack. Reynolds numbers are of the order of 1.0E5-3.0E5. Two different grid configurations, the coarse and the fine grids, are applied in order to speed up convergence. In coarse grid configuration, 92820 total grid points are used in two blocks, whereas it is 700,000 in fine grid. The flow field is dominated by the vortices and the separated flows. Baldwin Lomax turbulence model is used over the flat plate surface. For the regions dominated by the strong side edge vortices, turbulence model is modified using a polar coordinate system whose origin is at the minimum pressure point of the vortex. In addition, an algebraic wake-type turbulence model is used for the wake region behind the wing. The initial flow variables at the fine grid points are obtained by the interpolation based on the coarse grid results previously obtained for 40000 iterations. Iterations are continued with the fine grid about 20000-40000 more steps. Pressures of the top surface are predicted well with the exception of leading edge region, which may be due to unsuitable turbulence model and/or grid quality. The predictions of the side edge vortices and the size of the leading edge bubble are in good agreement with the experiment.

Flow separation is a source of aerodynamic in-efficiency; however by using vortex generators the issue of flow separation can be controlled. This is of particular benefit to flows around bluff bodies which are susceptible to large scale separated flows, such as bodies in ground effect. Previous studies concerning heat transfer applications focused on the ability of low aspect ratio spherical wells (dimples) to produce vortices for flow mixing. Dimpled surfacing on an inverted Tyrrell026 airfoil in ground effect (indicative of high performance automotive aerodynamic applications e.g. Formula One) has been investigated for similar vortex enhanced wake reductions. Experimental measurements using Laser Doppler Anemometry (LDA) and Particle Image Velocimetry (PIV) were taken inside a dimple to provide validation and verification of numerical analyses of dimple flows. The k-ω SST turbulence model showed good agreement to the experimental measurements. Additionally experiments were conducted using LDA and PIV with various configurations of dimple arrays placed from a fixed separation point of a 16˚ rearward facing ramp to determine how the array configuration influenced the large scale separation. The airfoil wake with numerous dimple configurations and placements were measured using LDA. Results showed that an array of dimples with close dimple to dimple spacing there was flow recovery in the airfoil wake from the velocity deficit with no dimples of u/Uo,min=-0.1, to u/Uo,min=0.4 with a dimple array, (at α=10˚, ground clearance h/c=0.313). At α=10˚ reductions in the wake size of 30%, 33%, 58% and 68% were found for the ground clearances of h/c=0.112, 0.134, 0.224 and 0.313 respectively. For numerous dimple array configurations, closely spaced dimple arrays were more effective in reducing the wake size, turbulence intensity and Reynolds stresses than those where dimple spacing was further apart. The chord wise location of the array on the wing affected the angle of incidence of the wing for which the wake was able to be reduced. Arrays placed towards the trailing edge improved wake losses at lower angles of incidence. Dimples placed further forward yielded the most improvement at higher angles of incidence, in part due to the increased venturi effect under the wing.

NR 20140805

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