Aerospace Engineering Theses and Dissertations

Permanent URI for this collectionhttp://hdl.handle.net/1903/2737

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    SCHLIEREN SEQUENCE ANALYSIS USING COMPUTER VISION
    (2013) Smith, Nathanial Timothy; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Computer vision-based methods are proposed for extraction and measurement of flow structures of interest in schlieren video. As schlieren data has increased with faster frame rates, we are faced with thousands of images to analyze. This presents an opportunity to study global flow structures over time that may not be evident from surface measurements. A degree of automation is desirable to extract flow structures and features to give information on their behavior through the sequence. Using an interdisciplinary approach, the analysis of large schlieren data is recast as a computer vision problem. The double-cone schlieren sequence is used as a testbed for the methodology; it is unique in that it contains 5,000 images, complex phenomena, and is feature rich. Oblique structures such as shock waves and shear layers are common in schlieren images. A vision-based methodology is used to provide an estimate of oblique structure angles through the unsteady sequence. The methodology has been applied to a complex flowfield with multiple shocks. A converged detection success rate between 94% and 97% for these structures is obtained. The modified curvature scale space is used to define features at salient points on shock contours. A challenge in developing methods for feature extraction in schlieren images is the reconciliation of existing techniques with features of interest to an aerodynamicist. Domain-specific knowledge of physics must therefore be incorporated into the definition and detec- tion phases. Known location and physically possible structure representations form a knowledge base that provides a unique feature definition and extraction. Model tip location and the motion of a shock intersection across several thousand frames are identified, localized, and tracked. Images are parsed into physically meaningful labels using segmentation. Using this representation, it is shown that in the double-cone flowfield, the dominant unsteady motion is associated with large scale random events within the aft-cone bow shock. Small scale organized motion is associated with the shock-separated flow on the fore-cone surface. We show that computer vision is a natural and useful extension to the evaluation of schlieren data, and that segmentation has the potential to permit new large scale measurements of flow motion.
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    THE EFFECTS OF FINITE-RATE REACTIONS AT THE GAS/SURFACE INTERFACE IN SUPPORT OF THERMAL PROTECTION SYSTEM DESIGN
    (2011) Beerman, Adam Farrell; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Gas-surface modeling is dependent on material type and atmospheric reentry conditions. Lower molecular collisions at the low pressure trajectories make it more likely for occurrences of nonequilibrium, or finite-rate, reactions. Equilibrium is often assumed at the surface of a material as it is a subset of nonequilibrium and is easier to compute, though it can lead to overly conservative predictions. A case where a low density material experiences a low pressure trajectory and designed for equilibrium is the Stardust Return Capsule (SRC) with the Phenolic Impregnated Carbon Ablator (PICA) as its heatshield. Post-flight analysis of the recession on the SRC found that the prediction from the equilibrium model can be more than 50% larger than the measured recession. The Modified Park Model was chosen as the finite-rate model as it contains simple four reactions (oxidation, sublimation, and nitridation) and has been previously used to study individual points of the SRC trajectory. The Modified Park Model cannot model equilibrium so a model BFIAT was developed that allows finite-rate reactions to be applied to the surface for a certain length of time. Finite-rate sublimation was determined to be reaction of importance in the Park Model for SRC-like conditions. The predicted recession on the SRC heatshield experienced a reduction in its overprediction; the finite-rate predictions fall with the measurement error of the recession at three points on the heatshield. The recession reduction was driven by a significant reduction in char formation. There was little change in the pyrolysis gas rate. The finite-rate model was also applied to simulations of various arc-jet tests that covered a range of heating conditions on the surface of the PICA material. Comparison to this experimental data further showed the role of finite-rate reactions and sublimation in the Park Model and conditions that favor the nonequilibrium assumption (heating over 1000 W/cm2). For the emerging PICA material, used for the Mars Science Laboratory and one of two material choices for the Crew Exploration Vehicle, and SRC-like trajectories, a finite-rate model was developed such that the more robust nonequilibrium assumption can be applied to design processes to reduce heatshield mass.
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    Computational Fluid Dynamic Solutions of Optimized Heat Shields Designed for Earth Entry
    (2010) Meeroff, Jamie Gabriel; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    Computational fluid dynamic solutions are obtained for heat shields optimized aerothermodynamically using Newtonian impact theory. Aerodynamically, the low-order approach matches computational simulations within 10%. Benchmark Apollo 4 solutions show that predicted heat fluxes under-predict convective heating by 30% and over-predict radiative heating by 16% compared to computational results. Parametric studies display a power law reliance of convective heat flux on edge radius. A slender heat shield optimized for a single design point produces heat fluxes 1.8 times what was predicted using the Newtonian approach. Here, maximum heating decreases with the inverse cube of the base sharpness. Coupled vehicle/trajectory optimized designs are examined for lunar return (11 km/s) and Mars return (12.5 km/s) and show possible discrepancies for eccentric shapes using low-order empirical correlations. Ultimately, gains suggested by the low-order approach using complex geometries are not reflected in high-fidelity simulations. In some respects, the simpler shape is the ideal one
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    Application of Compound Compressible Flow to Hypersonic Three-Dimensional Inlets
    (2009) Bussey, Gillian Mary Harding; Lewis, Mark J.; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    A method for correcting flow non-uniformities and incorporating multiple oblique shocks waves into compound compressible flow is presented. This method has several applications and is specifically presented for the problem of creating a streamline-traced hypersonic three-dimensional inlet. This method uses compound compressible flow theory to solve for the freestream flow entering a pre-defined duct with a desired downstream profile. This method allows for multiple iterations of the design space and is computational inexpensive. A method is also presented for modeling a laminar or turbulent boundary layer to compare inlet designs and to determine the viscous correction to the inlet. Two different Mach 6 designs were evaluated, with a rectangular capture area and circular combustor with a uniform temperature, pressure, and Mach number profile. Comparison with other three-dimensional inlets indicates those designed with this method demonstrate good inviscid performance. These inlets also have the ability to correct incoming flow non-uniformities.
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    Aerothermodynamic Optimization of Earth Entry Blunt Body Heat Shields for Lunar and Mars Return
    (2009) Johnson, Joshua E.; Lewis, Mark J; Aerospace Engineering; Digital Repository at the University of Maryland; University of Maryland (College Park, Md.)
    A 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.