Thesis (M.S.)--Worcester Polytechnic Institute. / Keywords: vaporizer; two-phase flow; electric propulsion. Includes bibliographical references (leaves 152-156).

Thesis (Ph. D.)--Aerospace Engineering, Georgia Institute of Technology, 2005. / Ben T. Zinn, Committee Chair ; Jeff Jagoda, Committee Member ; Suresh Menon, Committee Member ; Tim Lieuwen, Committee Member ; Rick Gaeta, Committee Member. Vita. Includes bibliographical references.

Thesis (Ph. D.)--University of Missouri-Columbia, 2008. / The entire dissertation/thesis text is included in the research.pdf file; the official abstract appears in the short.pdf file (which also appears in the research.pdf); a non-technical general description, or public abstract, appears in the public.pdf file. Title from title screen of research.pdf file (viewed on June 9, 2009) Vita. Includes bibliographical references.

Thesis (M.S.)--West Virginia University, 2005. / Title from document title page. Document formatted into pages; contains vii, 83 p. : ill. (some col.). Includes abstract. Includes bibliographical references (p. 53-57).

An investigation, both theoretical and experimental in nature, has been undertaken to develop a simple method for the design of supersonic nozzles and, indeed, inlets of quite complex shapes from known or calculated axisymmetric flowfields. 1he axisymmetric flowfield is determined using a computer program based on the method of characteristics. Streamlines are calculated by direct integration of the axisymmetric stream function. 7he desired shape is chosen at the exit of the computed axisymmetric nozzle having the desired length and Mach number. Its describing points are then traced along the corresponding streamlines back to the throat. Streamsheets formed by these streamlines define the new shape. Following this approach, two three-dimensional nozzles were designed : one of elliptical cross-section and a two-dimensional wedge. Flows within the two configurations were further simulated using a general purpose three-dimensional CFD code, 'PHOENICS', while the elliptical nozzle was subsequently manufactured and submitted to experimental tests. Results from the experimental tests and three-dimensional numerical simulation, as well as predictions of the performance of the nonaxisymmetric nozzles and their axisymmetric counterparts were obtained and compared. Good agreement was achieved between the several components of the study demonstrating that it is possible, using this relatively simple method, to design satisfactory three-dimensional nozzles.

The performance of an aircraft gas turbine engine is adversely affected by the non-uniform or distorted flow in the inlet duct. Inlet flow distortion lowers the surge margin of the engine‟s compression system with surge occurring at much lower pressure ratios at all engine speeds. The compressor and/or engine are subjected to ground tests in the presence of inlet distortion to evaluate its performance. The simplest method of simulating inlet distortion during these tests is by installing a distortion screen ahead of the engine on the test bed. The uniform inlet flow to the compressor becomes nonuniform with total pressure loss after passing through the distortion screen. Though the distortion screens offer a number of significant advantages, they have some disadvantages. The air jet distortion system can alleviate many of the operational disadvantages encountered with the conventional distortion screens. The system consists of a number of air jets arranged in a circumferential array in a plane and issuing opposite to the primary air flow entering the engine. The jets interact with the primary stream and cause a local total pressure loss due to momentum exchange. The individual mass flow rates from the jets can be varied to obtain a required total pressure pattern ahead of the compressor at the Aerodynamic Interface Plane (AIP). A systematic study of the flow field of confined, turbulent, incompressible, axisymmetric jet issuing into counterflow is covered in this research programme. The jet penetration length and the jet width are reduced compared to unconfined counterflow and a linear relationship between the velocity ratio and the jet length ceases to be valid. The flow field of a circular compressible turbulent jet and then a system of four jets arranged circumferentially and issuing into a confined counterflow was studied experimentally and numerically. For the four jet system the mass flow rates in the four jets were equal in the first part of the study and in the second part they were unequal. The loss in total pressure due to the jet(s) interacting with the counterflow was quantified by a total pressure loss parameter λp0. The total pressure loss increased with increasing mass flow ratio. The total pressure loss distribution was evaluated at several locations behind the jet injector(s). The total pressure non-uniformity quantified by Distortion Index (DI) was found to be highest at a location just downstream of the jet injector and at far downstream locations low values of DI were observed. From the understanding gained with a single jet and four jets in counterflow a methodology was developed to generate a given total pressure distortion pattern at the AIP. The methodology employs computations to obtain the total pressure distortion at the AIP with quasi-one-dimensional inviscid analysis used as a starting point to estimate the mass flow rate in the jets. The inviscid analysis also provides a direction to the iterative procedure to vary the mass flow rate in the jets at the end of each computational step. The methodology is demonstrated to generate a given total pressure distortion pattern using four jets and is further extended to a larger number of jets, twelve and later twenty jets. The total pressure distortion patterns typical of use in aircraft gas turbine engine testing are generated accurately with a smaller number of jets than reported in the literature.

Applied research and testing has been conducted at the Cal Poly San Luis Obispo High-pressure Blow-Down facility to study the affects of throttling in a thrust-vectored aerospike nozzle. This study supports the ongoing research at Cal Poly to effectively thrust vector a hybrid rocket motor. Such thrust vectoring is achieved by small secondary ports in the nozzle body that are perpendicular to the main nozzle. The testing conducted included characterizing and comparing the performance of a straight aerospike nozzle to that of a thrust-vectored aerospike nozzle. Throttling effects on the aerospike nozzle in an unvectored and in a vectored configuration were also investigated. The interaction between throttling and thrust vectoring of an aerospike nozzle is the focus of this thesis research. This research shows that large-throat/high-thrust operation of an aerospike nozzle provides little thrust vector generation. Conversely, small-throat/low-thrust operation provides ample thrust vector generation. These results have implications in the effectiveness of thrust vectoring an aerospike nozzle with secondary ports. Rockets having an aerospike nozzle with throttling capabilities will be subject to the minimum and maximum turn angles for a given throttle position. As such, certain vehicle maneuvers might not be obtainable at certain throttle operations. Conversely, at lower throttling conditions, higher turn angles will be achievable.

In this work, laser ablation is investigated as a possible propulsion technique for space applications.

The objective of this investigation was to improve the thermal environment for a turbine vane through reduction of passage secondary flows. This was accomplished by modifying the vane/endwall junction to include a leading edge fillet. The problem approach was to integrate optimization methods with computational fluid dynamics to optimize the fillet design. The resulting leading edge fillet was then tested in a large-scale, low speed cascade to verify thermal performance. A combustor simulator located upstream of the cascade was used to generate realistic inlet conditions for the turbine vane. Both computational and experimental results underscore the importance of properly modeling the inlet conditions to the turbine. Results of the computational optimization process indicate that significant reductions in adiabatic wall temperature can be achieved with a leading edge fillet. While the intent of the initial fillet design was to improve the thermal environment for the vane endwall, computational results also indicate thermal benefit to the vane surfaces. Flow and thermal field results show that a fillet can enhance coolant effectiveness, prevent formation of the leading edge horseshoe vortex, and preclude full development of a passage vortex. In experimental testing, four cascade inlet conditions were investigated to evaluate the effectiveness of the fillet in reducing endwall temperature levels. Two tested conditions featured a flush combustor/cascade interface, while the remaining two included coolant injection through a backward-facing slot. With the flush interface, fillet thermal performance was evaluated for two inlet total pressure profiles. For the design profile, the fillet had a positive impact on the endwall temperature distribution as well as on the passage thermal field. For the off-design profile, the fillet was observed to have a slightly detrimental impact on the endwall adiabatic temperature distribution; however, passage thermal field results indicate a thermal benefit for the vane suction surface. With the backward-facing slot, thermal tests were conducted for two slot coolant flow rates. For both slot flow rates, the fillet improved endwall thermal protection and prevented coolant lift-off. While increasing the flow rate of slot coolant enhanced endwall effectiveness, fillet thermal performance was similar for the two slot flow rates. / Ph. D.

Small-scale marine propellers operate at low Reynolds numbers, where laminar-turbulent transition of the boundary layer can impact the distributions of pressure and shear stress on the blade surface. Marine propellers operating at low Reynolds numbers are subject to laminar-turbulent transition of the boundary layer, which impacts the distributions of pressure and shear stress on the blade surface. To design efficient propellers for operation at low Reynolds numbers, transitional effects must be included in the evaluations of propeller performance. In this work, transition modeling techniques in Reynolds Averaged Navier-Stokes computational fluid dynamics (RANS CFD) are utilized to evaluate and design propellers operating at low Reynolds numbers. The Galilean invariant γ transition model with an extension for crossflow transition is coupled to the SSG (Speziale, Sarkar, Gatski) /LRR (Launder, Reece, Rodi) -ω Reynolds stress transport turbulence model, with validation cases performed for flate plate boundary layers, 2-dimensional airfoils, a 3-dimensional wing, and 6:1 prolate spheroids. The performance of the coupled SSG/LRR-ω-γ Reynolds stress transition model for propellers with flow transition is then evaluated using experimental surface streamline and force data from four model-scale marine propellers. A method for the design of low Reynolds number marine propellers is presented using a transition-sensitive lifting line method coupled with the panel method code XFOIL. Initial geometries generated using the lifting-line method are then optimized in RANS CFD using the 2 equation γ-Reθ transition model and an adjoint method to warp the propeller shape to improve the efficiency. Two design studies are performed, including an open water propeller, and a propeller designed for a small autonomous underwater vehicle. / Doctor of Philosophy / Small-scale marine propellers exhibit transition from laminar to turbulent flow in the region near the surface of the blades. Regions of laminar and turbulent flow on the blade surface contribute differently to the overall thrust and torque on the propeller. Prediction of flow transition in the design process for small-scale marine propellers can improve the accuracy of the thrust and torque prediction compared to modeling the flow as purely laminar or turbulent. Propeller thrust and torque can be modeled using computational fluid dynamics (CFD) simulations, where transition modeling is accomplished by solving a transport equation for the intermittency γ, which represents the percentage of time the flow in a given location is turbulent. In this work, a transition model is coupled to a high-fidelity full Reynolds stress turbulence model, which solves 6 transport equations to solve for each component of the Reynolds stress tensor. The Reynolds stress tensor represents the turbulent velocity fluctuations in the governing equations solved in the CFD simulation. This coupled transition and turbulence model is then validated using experimental results of flows with a number of different transition mechanisms. The coupled model is then tested with a series of model-scale propellers, with results of the CFD simulations compared to the experimental results. A method for the design of propellers with flow transition is presented which incorporates transition effects. The designs generated by this method are then optimized in a CFD framework which morphs the blade geometry to improve the ratio of the thrust produced by the propeller to the torque, which corresponds to a higher efficiency. Two design cases are presented: a propeller designed for open water operation, and a propeller design for a small autonomous underwater vehicle.