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dc.contributor.advisorLewis, Mark Jen_US
dc.contributor.authorJohnson, Joshua E.en_US
dc.date.accessioned2009-07-02T05:47:40Z
dc.date.available2009-07-02T05:47:40Z
dc.date.issued2009en_US
dc.identifier.urihttp://hdl.handle.net/1903/9177
dc.description.abstractA differential evolutionary algorithm has been executed to optimize the hypersonic aerodynamic and stagnation-point heat transfer performance of Earth entry heat shields for Lunar and Mars return manned missions with entry velocities of 11 and 12.5 km/s respectively. The aerothermodynamic performance of heat shield geometries with lift-to-drag ratios up to 1.0 is studied. Each considered heat shield geometry is composed of an axial profile tailored to fit a base cross section. Axial profiles consist of spherical segments, spherically blunted cones, and power laws. Heat shield cross sections include oblate and prolate ellipses, rounded-edge parallelograms, and blendings of the two. Aerothermodynamic models are based on modified Newtonian impact theory with semi-empirical correlations for convection and radiation. Multi-objective function optimization is performed to determine optimal trade-offs between performance parameters. Objective functions consist of minimizing heat load and heat flux and maximizing down range and cross range. Results indicate that skipping trajectories allow for vehicles with L/D = 0.3, 0.5, and 1.0 at lunar return flight conditions to produce maximum cross ranges of 950, 1500, and 3000 km respectively before Qs,tot increases dramatically. Maximum cross range increases by ~20% with an increase in entry velocity from 11 to 12.5 km/s. Optimal configurations for all three lift-to-drag ratios produce down ranges up to approximately 26,000 km for both lunar and Mars return. Assuming a 10,000 kg mass and L/D = 0.27, the current Orion configuration is projected to experience a heat load of approximately 68 kJ/cm2 for Mars return flight conditions. For both L/D = 0.3 and 0.5, a 30% increase in entry vehicle mass from 10,000 kg produces a 20-30% increase in Qs,tot. For a given L/D, highly-eccentric heat shields do not produce greater cross range or down range. With a 5 g deceleration limit and L/D = 0.3, a highly oblate cross section with an eccentricity of 0.968 produces a 35% reduction in heat load over designs with zero eccentricity due to the eccentric heat shield's greater drag area that allows the vehicle to decelerate higher in the atmosphere. In this case, the heat shield's drag area is traded off with volumetric efficiency while fulfilling the given set of mission requirements. Additionally, the high radius-of-curvature of the spherical segment axial profile provides the best combination of heat transfer and aerodynamic performance for both entry velocities and a 5 g deceleration limit.en_US
dc.format.extent31751078 bytes
dc.format.mimetypeapplication/pdf
dc.language.isoen_US
dc.titleAerothermodynamic Optimization of Earth Entry Blunt Body Heat Shields for Lunar and Mars Returnen_US
dc.typeDissertationen_US
dc.contributor.publisherDigital Repository at the University of Marylanden_US
dc.contributor.publisherUniversity of Maryland (College Park, Md.)en_US
dc.contributor.departmentAerospace Engineeringen_US
dc.subject.pqcontrolledEngineering, Aerospaceen_US
dc.subject.pquncontrolledaerothermodynamicsen_US
dc.subject.pquncontrolledentry vehicleen_US
dc.subject.pquncontrolledheat shielden_US
dc.subject.pquncontrolledhypersonicsen_US
dc.subject.pquncontrolledlunar and mars returnen_US
dc.subject.pquncontrolledreentryen_US


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