An Axisymmetric Rocket Ejector Simulation (ARES) was developed to numerically analyze various configurations of an air augmented rocket. Primary and secondary flow field visualizations are presented and performance predictions are tabulated. A parametric study on ejector geometry is obtained following a validation of the flow fields and performance values. The primary flow is calculated using a quasi-2D, irrotational Method of Characteristics and the secondary flow is found using isentropic relations. Primary calculations begin at the throat and extend through the nozzle to the location of the first Mach Disk. Combustion properties are tabulated before analysis to allow for propellant property selection. Secondary flow calculations employ the previously calculated plume boundary and ejector geometry to form an isentropic solution. Primary and secondary flow computations are iterated along the new pressure distributions established by the 1D analysis until a convergence tolerance is met. Thrust augmentation and Specific Impulse values are predicted using a control volume approach. For the validation test cases, the nozzle characteristic net is very similar to that of previous research. Plume characteristics are in good agreement but fluctuate in accuracy due to flow structure formulation. The individual unit processes utilized by the Method of Characteristics are found to vary their outputs by up to 0.025% when compared to existing sources. Rocket thrust and specific impulse are increased by up to 22% for a static system and 15% for an ejector flow at Mach 0.5. Evidence of Fabri conditions were observed in the flow visualization and graphically through the performance predictions. It was determined that the optimum ejector divergence angle for an air augmented rocket greatly depends on the stagnation pressure ratio between the primary and secondary flows.

A transformation between minimum dimension adjoint variables and redundant adjoint variables is derived in this dissertation. The transformation is then applied between the adjoint variables associated with Cartesian position and velocity vectors and a set of redundant adjoint variables associated with certain regularized variables (Schumacher variables). This transformation proves to be very beneficial when it is applied to minimum-fuel space rendezvous and intercept problems. It facilitates using attributes from the two systems simultaneously; a new necessary condition in Schumacher adjoints is derived in this dissertation, and this together with classical necessary conditions for fuel-optimal transfer (existing in the position and velocity space) leads to a numerical algorithm which seems to be quite robust in finding candidate optimal control solutions for space transfer problems. / Ph. D.

Turbine blade surface temperatures were studied during transient operation in a turbofan engine test rig. A single fiber radiation pyrometer was used to view the suction side of the blades from approximately 60 percent axial chord to the trailing edge at an average radial location of 70 percent blade height. A single ceramic-coated blade produced a once-per-revolution signal that allowed for the tracking of individual blades during the transients. The investigation concentrated on the light-off starting transient and the transients obtained during accelerating and decelerating between power settings. During starting and acceleration transients, the blade surface temperature gradient was observed to reverse. This phenomenon was most apparent during starting when the trailing edge was initially much hotter than the 60 percent chord location, resulting in large temperature gradients. In steady operation the trailing edge temperature was lower than the 60 percent chord location, and the gradients were less severe. During deceleration transients, the trailing edge cooled more rapidly than the 60 percent chord location. This resulted in larger temperature gradients than were seen in steady operation, but no profile inversion was observed. These temperature gradients and profile inversions represent a cycling of thermally-induced stresses which may contribute to low cycle fatigue damage. A simple one-dimensional heat transfer model is presented as a means of explaining the different heating rates observed during the transients. / Master of Science

One-Dimensional modeling of dual mode scramjet and ramjet flowpaths is a useful tool for scramjet conceptual design and wind tunnel testing. In this thesis, modeling tools that enable detailed analysis of the flow physics within the combustor are developed as part of a new one-dimensional MATLAB-based model named VTMODEL. VTMODEL divides a ramjet or scramjet flow path into four major components: inlet, isolator, combustor, and nozzle. The inlet module provides two options for supersonic inlet one-dimensional calculations; a correlation from MIL Spec 5007D, and a kinetic energy efficiency correlation. The kinetic energy efficiency correlation also enables the user to account for inlet heat transfer using a total temperature term in the equation for pressure recovery. The isolator model also provides two options for calculating the pressure rise and the isolator shock train. The first model is a combined Fanno flow and oblique shock system. The second model is a rectangular shock train correlation. The combustor module has two options for the user in regards to combustion calculations. The first option is an equilibrium calculation with a "growing combustion sphere" combustion efficiency model, which can be used with any fuel. The second option is a non-equilibrium reduced-order hydrogen calculation which involves a mixing correlation based on Mach number and distance from the fuel injectors. This model is only usable for analysis of combustion with hydrogen fuel. Using the combustion reaction models, the combustor flow model calculates changes in Mach number and flow properties due to the combustion process and area change, using an influence coefficient method. This method also can take into account heat transfer, change in specific heat ratio, change in enthalpy, and other thermodynamic properties. The thesis provides a description of the flow models that were assembled to create VTMODEL. In calculated examples, flow predictions from VTMODEL were compared with experimental data obtained in the University of Virginia supersonic combustion wind tunnel, and with reported results from the scramjet models SSCREAM and RJPA. Results compared well with the experiment and models, and showed the capabilities provided by VTMODEL. / Master of Science

”The Eagle has landed”. With this words, astronaut Neil Armstrong, together with astronaut Edwin ”Buzz” Aldrin, confirmed the safe landing on the Moon of the Apollo11 on July 20th, 1969. This achievement would have never been possible without the engineering behind the construction of the Moon lander. One of the most innovative features installed aboard was the pintle injector, a specific type of injector capable of precise metering of the propellants into the combustion chamber, and capable of throttling. Furthermore, the pintle injector has demonstrated its inherent combustion stability andgood mixing properties over time. The current work serves as a feasibility study for the use of an electric motor as the actuator for the pintle. This paper is the result of a 9-month internship at ArianeGroup GmbH, which is investigating the use of a pintle injector as a back-up option for a future ESA lunar lander mission. A preliminary design of the pintle was already produced, and the scope of this work is the design of the mechanical linkages and the actuation of the injector, with the choice of the method of actuation, the electric motor and the necessary components.

Pressure gradient confinement effects are experimentally investigated within a cavity combustor to analyze the flame interactions of premixed, cavity stabilized, flames in a high-speed combustor. Pressure gradient confinement effects are generated in a dual mode ramjet-scramjet (DMSR) by varying the wall geometry to form converging, diverging, and nominal configurations. The velocity field and flame position are captured temporally using simultaneous high-speed particle image velocimetry (PIV) and CH chemiluminescence. The evolution of the flow field and flame structure are analyzed, and the high temporal resolution of these measurements allows for the characterization of turbulence-flame interactions. Consideration of the combustion mode and inlet conditions, such as the inflow velocity and turbulence, are vital in studying flame-vortex interaction dynamics and its effect on the flame stabilization process and is essential in ensuring efficient, complete combustion. The results from the experiment will provide a greater understanding of how flame-vorticity interactions and pressure gradient confinement effects play a key role in the flame-stabilization and combustion process.

No description available.

The contributions of vortex stretching, dilatation, baroclinic torque, and viscous diffusion to vorticity transport are experimentally investigated in a high-Reynolds number cavity combustor using high-speed particle image velocimetry and broadband chemiluminescence. An adaptive wall geometry forming converging, diverging, and nominal configurations is implemented to study the effects of pressure gradient on local flow physics and vorticity dynamics. The spatial profiles of the local turbulence terms are conditioned on the mean flame front to characterize the influence of the pressure gradient field and exothermic heat release on vortex dynamics in the cavity. In addition to isolate the influence of combustion on the flow, a nonreacting analysis is performed and a correlation is made between combustor geometry and the turbulence transport processes. Vorticity transport through dilatation was found to be significant relative to the other transport terms across all the configurations studied. These results contrast with direct numerical simulations of high Reynolds number flows in homogeneous isentropic turbulence. In addition, a scaling is proposed to quantify the significance of the flow induced vorticity and pressure fields on dilatation and baroclinic torque vorticity production. Experimental studies of similar confined combustors show a similar trend to the numerical studies with baroclinic torque dominating the transport mechanisms, motivating this study to understand the dependence of vorticity transport on the underlying flow physics.

Electrically controllable solid propellants are an area of interest as a viable solution to the lack of throttle-ability in solid propellant rocket motors. Existing studies have focused on propellants compositions using hydroxyl-ammonium nitrate, ammonium nitrate, or lithium perchlorate as oxidizers. Additionally, the thermochemical and electrochemical reaction mechanisms have not yet been fully defined. The research in this thesis explores the nitrate and perchlorate oxidizer families to compare their cation-anion relationships. Using these oxidizers, pseudo electrically controllable solid propellant compositions were created with the addition of multi-wall carbon nanotubes to enhance ohmic heating capabilities. These additives were selected based on theory that with a non-complexing polymer, an oxidizer melt layer is required for ions to dissociate and electrically controlled ignition to occur. Using an applied voltage, ignition delay and current draw experiments were performed to expand on prior findings that ignition delay follows oxidizer melt temperature while mobility is associated with the size of the ionic radii. Additionally, neat oxidizer pellets were electrically decomposed to determine their linear regression rate. These results help to characterize the mechanism of reaction. This advances the knowledge of oxidizers in electrically controllable applications. / Master of Science / Solid propellant rocket motors have been extensively studied and used in both space and military applications because they do not use air as the source of oxygen. Their main limitation is the lack of throttle-ability, or inability to control propellant burning. This is because solid propellants, which are generally composed of an ionic oxidizer salt, a polymer fuel, and additives, are pre-combined and stored within the rocket motor. An emerging viable solution to this limitation is electrically controllable solid propellants. With an applied voltage, the oxidizer is heated and melts, allowing ions to dissociate and current to flow between electrodes. This reaction can then be controlled by turning the power supply on and off. Cations, or ions which have a net positive charge, move to the negatively charged cathode while anions, which have a net negative charge, move to the negatively charged anode. The research in this thesis explores different cation-anion oxidizer pairings using both a propellant composition and as a pure oxidizer under an applied voltage. The results help to characterize the mechanism of reaction of each oxidizer in an electrically controllable context and determine their effectiveness in these propellant applications.

The design of aircraft propulsion configurations must digress from the typical configurations that are utilized on the majority of aircraft in order to consider the effects of environmental issues as well as the noise that is generated from the engines. One unconventional approach under consideration involves rectangular jets near flat surfaces that are parallel to the jet axis. This type of configuration makes an attempt to muffle the noise that propagates to the ground, but previous experimental work showed that the noise generated by this configuration was actually increased due to the effect that the plate trailing edge exerts on the flow. In this thesis, a large eddy simulation study is conducted to determine whether wall deformations at the plate trailing edge could reduce the jet noise. A high aspect ratio rectangular nozzle is placed over a flat surface featuring sinusoidal deformations at the trailing edge. A range of amplitudes and wavenumbers, characterizing the deformations at the trailing edge, is considered to determine the parameter range that corresponds to noise reduction.