Browsing by Author "Estruch-Samper, David"
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Item Open Access A case study on the aerodynamic heating of a hypersonic vehicle(Royal Aeronautical Society, 2012-09-30T00:00:00Z) Mifsud, Michael; Estruch-Samper, David; MacManus, David G.; Chaplin, Ross; Stollery, J. L.A Parabolised Navier-Stokes (PNS) flow solver is used to predict the aerodynamic heating on the surface of a hypersonic vehicle. This case study highlights some of the main heat flux sensitivies to various conditions for a full-scale vehicle and illustrates the use of different complimentary methods in assessing the heat load for a realistic application. Different flight phases of the vehicle are considered, with freestream conditions from Mach 4 to Mach 8 across a range of altitudes. Both laminar and turbulent flows are studied, together with the effect of the isothermal wall temperature, boundary-layer transition location and body incidence. The effect of the Spalart-Allmaras and Baldwin-Lomax turbulent models on the heat transfer distributions is assessed. A rigorous assessment of the computations is conducted through both iterative and grid convergence studies and a supporting experimental investigation is performed on a 1/20th scale model of the vehicle's forebody for the validation of the numerical results. Good agreement is found between the PNS predictions, measurements and empirical methods for the vehicle forebody. The present PNS approach is shown to provide useful predictions of the heat transfer over the axisymmetric vehicle body. A highly complex flow field is predicted in the fin-body-fin region at the rear of the vehicle characterised by strong interference effects which limit the predictions over this region to a predominately qualitative level.Item Open Access Hypersonic interference aerothermodynamics(2009-10) Estruch-Samper, David; Lawson, Nicholas J.; Garry, Kevin P.When a vehicle travels at hypersonic speeds during launch, cruise or atmospheric re-entry it is subject to extremely high surface flow temperatures. As well as on the vehicle forebody, extreme heating can take place close to surface protuberances which are almost impossible to avoid in a real flight vehicle. These disturbances interfere with the freestream flow and result in complex viscous interactions which induce a local heat flux augmentation that can become detrimental to the integrity of the vehicle. A greater understanding of these flow phenomena is required. This thesis develops the understanding of the behaviour of the flow around surface protuberances in hypersonic vehicles and presents an engineering approach to predict the location and magnitude of the highest heat transfer rates in their vicinity. To this end, an experimental investigation was performed in a gun tunnel at freestream Mach numbers of 8.2 and 12.3 and Reynolds numbers ranging from Reoo/m=3.35xl0 ⁶ to Reꚙ /m=9.35xl0 ⁶. The effects of protuberance geometry, boundary layer state, freestream Reynolds number and freestream Mach number were assessed. Further understanding of the flowfield was obtained through oil-dot visualisations and highspeed schlieren videos taken at frame rates of up to 50 kHz. Results show the local interference interaction is strongly three-dimensional and is dominated by the incipient separation angle induced by the protuberance. In subcritical interactions - in which the incoming boundary layer remains unseparated upstream of the protuberance - the highest heating occurs adjacent to the device. In supercritical interactions - in which the incoming boundary layer is fully separated ahead of the protuberance - the highest heating generally occurs on the surface just upstream of it. An exception is for low-deflection protuberances under low-Reynolds freestream flow conditions in which case the heat flux to the side is greater.Item Open Access Hypersonic interference heating in the vicinity of surface protuberances(Springer Science Business Media, 2010-09-30T00:00:00Z) Estruch-Samper, David; MacManus, David G.; Stollery, J. L.; Lawson, Nicholas J.; Garry, Kevin P.The understanding of the behaviour of the flow around surface protuberances in hypersonic vehicles is developed and an engineering approach to predict the location and magnitude of the highest heat transfer rates in their vicinity is presented. To this end, an experimental investigation was performed in a hypersonic facility at freestream Mach numbers of 8.2 and 12.3 and Reynolds numbers ranging from Re (a)/m = 3.35 x 10(6) to Re (a)/m = 9.35 x 10(6). The effects of protuberance geometry, boundary layer state, freestream Reynolds number and freestream Mach numbers were assessed based on thin-film heat transfer measurements. Further understanding of the flowfield was obtained through oil-dot visualizations and high-speed schlieren videos. The local interference interaction was shown to be strongly 3-D and to be dominated by the incipient separation angle induced by the protuberance. In interactions in which the incoming boundary layer remains unseparated upstream of the protuberance, the highest heating occurs adjacent to the device. In interactions in which the incoming boundary layer is fully separated ahead of the protuberance, the highest heating generally occurs on the surface just upstream of it except for low-deflection protuberances under low Reynolds freestream flow conditions in which case the heat flux to the side is greater.Item Open Access Measurement of shock wave unsteadiness using a high-speed schlieren system and digital image processing(American Institute of Physics, 2008-12) Estruch-Samper, David; Lawson, Nicholas J.; MacManus, David G.; Garry, Kevin P.; Stollery, J. L.A new method to measure shock wave unsteadiness is presented. Time-resolved visualizations of the flow field under investigation are obtained using a high-speed schlieren optical system and the motion of the shock wave is determined by means of digital image processing. Information on the shock’s unsteadiness is subsequently derived with Fourier analysis. A sample study on shock unsteadiness in a shock-wave/turbulent boundary-layer interaction with separation is included. The method presented enables a measure of shock unsteadiness at locations in the imaged flow field not accessible by intrusive methods.