Thesis (Ph. D.)--University of Wisconsin--Madison, 1983. / Typescript. Vita. eContent provider-neutral record in process. Description based on print version record. Includes bibliographical references (leaves 77-78).

Thesis (Ph. D.)--University of Wisconsin--Madison, 1983. / Typescript. Vita. eContent provider-neutral record in process. Description based on print version record. Includes bibliographical references.

Thesis (Ph. D.)--University of Wisconsin--Madison, 1983. / Typescript. Vita. eContent provider-neutral record in process. Description based on print version record. Includes bibliographical references.

Thesis (Ph.D.)--Boston University. / An analysis is given which treats experimental heat-transfer data obtained during flights of three V-2 rockets and two Aerobee rockets. This work was conducted largely at Boston University under Air Force Contracts W28-099-ac-395 and AF 33(615)-382. The instrumentation and rocket-flight phases extended from 1947 to 1952, while the analysis was carried out in the period from 1953 to 1955. A total of nine V-2 rockets were instrumented to some extent for the purposes of this experiment. The missiles for which data have been analyzed were chosen on the basis of best performance on the part of both instrumentation and rocket. Experimental equipment of special design was produced and installed by the Boston University group in the rockets which were launched in the White Sands area in New Mexico. Extensive field trips were required in this connection, both for the experiment on heat0transer and others in which the same group participated at the same time. The general program was one of research in physics of the upper atmosphere. Since the conclusion of the supporting contracts in October of 1954, additional consideration of the experimental results has continued on a private basis. The data obtained during the course of the present experiment were in the form of temperature variations as experienced by heat meters which were placed at selected locations over the body and fin of the rockets. The heat meters were small circular diaphragms of stainless steel which were placed flush with the rocket skin. Each diaphragm had on its central inner surface a pair of resistance thermometers made of nickel wire, with a diameter of 0.002". The thermometers were switched in sequence into a resistance bridge. One thermometer had always a higher resistance than the other, an arrangement which served two purposes. Firstly, the two values of voltage for a given temperature permited unambiguous determination of the direction in which the resistance bridge was unbalanced. Secondly, these two values of voltage were handled independently by our own instrumentation, and by the recording or telemetry devices. When two such different values of voltage were finally translated back into their equivalent temperatures, close agreement of the temperature values lent considerable support to their accuracy. When the two thermometers reported temperatures which were significantly different from each other, restraint was indicated in trusting their accuracy. Fortunately, close agreement of temperatures reported by the dual thermometers was the more usual experience. [TRUNCATED]

Thesis (M.S.)--Virginia Polytechnic Institute and State University, 1991. / Vita. Abstract. Includes bibliographical references (leaves 122-125). Also available via the Internet.

The present thesis project consisted of two parts. First, a general method for determining the transient skin temperatures of bodies during high-speed flight was developed. The governing differential equation was presented for this purpose, giving the fundamental relations between the transient skin temperature and flight history. The determination of all pertinent parameters in the equation was discussed, and the Runge-Kutta numerical method of integration was used to obtain the solution. The method was employed to compute the time history of the skin temperatures for several hypothetical flight plans, and the results presented in the form of graphs. For the Mach number and altitude range investigated, the maximum skin temperature obtained was approximately 2200 °R and was found to be largely independent of the type of trajectory. The second portion of the project consisted of determining the effect of heat insulation on the cooling requirements of the internal structure of a high-speed vehicle. The governing equation for heat conduction through an isotropic solid was developed, and then modified to account for nonhomogeneous materials. The initial and boundary conditions for the governing equation were specified, and the equation solved by the method of finite-differences. The temperatures obtained, the first portion the thesis, were used as the outer surface temperature variation of the insulation, and the time history of the inner surface temperature of the insulation (for several thicknesses) was calculated. To make the problem as general as possible, the results were presented in terms of the thermal diffusivity of the insulating material. For illustrative purposes, an example problem was worked using rock wool as the insulating material. It was found that, by using one-half inch of this insulating material, the maximum temperature obtained by the internal structure was less that 5 percent of the skin temperature. Thus, it was concluded that the increase of the temperature of the internal structure of a high-speed vehicle during a limited time of flight, can be held to structurally permissable values by the use of heat insulation placed between the skin and the internal structure of the vehicle. / Master of Science

Thesis (Ph.D.)--Duquesne University, 2006. / Title from document title page. Abstract included in electronic submission form. Includes bibliographical references and abstracts.

Due to renewed interest in the sub orbital rocket program of the Florida Space Authority and a surplus of Super Loki Sounding Rockets, an effort to improve the usefulness of this surplus is herein undertaken. Currently, the capacity of the payload compartment in the upper stage of the Super Loki system is very limited. A redesign of the upper stage will allow larger and more versatile payloads to be carried, assuming the appropriate design parameters are met. It has therefore been undertaken to create a design procedure that is comprehensive in scope in order to affect this redesign. This procedure includes five major components. These are the separation of the upper and lower stages, the stability of the vehicle, the altitude and velocity of the rocket, the mechanical loading and finally the aerodynamic heating. Semi-empirical methods were used whenever possible to allow comparison with experimental data. This procedure revealed that larger diameter upper stages might be used up to a reasonable maximum of four inches. The four-inch modification is found to be stable as were the smaller modifications considered. The altitude and velocity of the rocket were found via a simple Eulerian time stepping scheme resulting in an estimate of approximately 148,000ft for the four-inch dart. The mechanical loading analysis allowed for the material selection for the rocket components. Reinforced steel fins and carbon fiber tubing, for the payload section, are adequate to meet expected mechanical loads, those being, 16000psi for the fin section due to launcher forces, 22800psi for compressive plus torsion forces on the composite section and 18000psi for the ejection stresses. An ablative coating is considered necessary to counteract the 760ºF temperatures along the composite tube.

No description available.

An approximate scheme for the rapid calculation of first estimates to the laminar heat transfer distribution over isothermal axisymmetric blunt bodies is developed. The method devised is free of any integral relations and reduces the required computing effort to a simple slide rule task. The simplicity of the method is due to the introduction of a new heat transfer parameter which is shown, from a semiempirical study, to undergo only moderate variation in regions where the heat transfer experiences order of magnitude changes. Based on these results, a series expansion for the parameter of interest is obtained through the fourth order term. Even though the perfect gas laws are employed in the series expansion, the resulting effect on the heat transfer ratio is felt to·be small. To substantiate the method, the heat transfer computed by the present scheme was compared with experimental, first-order exact, and Lees' approximate scheme for six body shapes of general interest. In all cases, fair to moderately good results were obtained. It is felt that any loss in accuracy is readily compensated for by the fact that the present method requires no numerical integration and therefore is extremely easy to apply. / Master of Science