Thesis. 1976. Sc.D.--Massachusetts Institute of Technology. Dept. of Aeronautics and Astronautics. / Microfiche copy available in Archives and Barker. / Vita. / Includes bibliographical references. / Sc.D.

No description available.

First, we investigate the steady Euler flows through a general 3-D axially symmetric infinitely long nozzles without irrotationality. Global existence and uniqueness of subsonic solution are established, when the variation of Bernoulli's function in the upstream is sufficiently small and mass flux has an upper critical value. / Second, we concerns the following transonic shock phenomena in a class of de Laval nozzles with porous medium posed by Courant-Friedrichs: Given a appropriately large receiver pressure pr, if the upstream flow is still supersonic behind the throat of the nozzle, then at a certain place in the diverging part of the nozzle a shock front intervenes and the gas is compressed and slowed down to subsonic speed. The position and the strength of the shock front are automatically adjusted so that the end pressure at the exit becomes pr. We investigate this problem for the full Euler equations, the stability of the transonic shock is proved when the upstream supersonic flow is a small steady perturbation of the uniform supersonic flow and the corresponding pressure at the exit has a small perturbation. / Duan, Ben. / Adviser: Zhouping Xin. / Source: Dissertation Abstracts International, Volume: 73-01, Section: B, page: . / Thesis (Ph.D.)--Chinese University of Hong Kong, 2010. / Includes bibliographical references (leaves 125-137). / Electronic reproduction. Hong Kong : Chinese University of Hong Kong, [2012] System requirements: Adobe Acrobat Reader. Available via World Wide Web. / Electronic reproduction. [Ann Arbor, MI] : ProQuest Information and Learning, [201-] System requirements: Adobe Acrobat Reader. Available via World Wide Web. / Abstract also in Chinese.

Thesis (Ph.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1979. / MICROFICHE COPY AVAILABLE IN ARCHIVES AND AERONAUTICS. / Vita. / Includes bibliographical references. / by Woodrow Whitlow, Jr. / Ph.D.

Weng, Shangkun. / Thesis (M.Phil.)--Chinese University of Hong Kong, 2009. / Includes bibliographical references (p. 73-78). / Abstracts in English and Chinese. / Abstract --- p.i / Acknowledgement --- p.iii / Chapter 1 --- Introduction --- p.1 / Chapter 2 --- Existence and Uniqueness results of transonic shock solution to full Euler system in a large variable nozzle --- p.11 / Chapter 2.1 --- The mathematical description of the transonic shock problem and main results --- p.11 / Chapter 2.2 --- The reformulation on problem (2.1.1) with (2.1.5)-(2.1.9) --- p.18 / Chapter 2.3 --- An Iteration Scheme --- p.30 / Chapter 2.4 --- A priori estimates and proofs of Theorem 2.2.1 and Theorem 2.1.1 --- p.39 / Chapter 3 --- A monotonic theorem on the shock position with respect to the exit pressure --- p.50 / Chapter 4 --- Discussions and Future work --- p.64 / Chapter 5 --- Appendix --- p.66 / Chapter 5.1 --- Appendix A: Background solution --- p.66 / Chapter 5.2 --- Appendix B: An outline of the proof of Theorem 2.1.2 --- p.67

Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1997. / Vita. Includes bibliographical references (p. 115-121).

In this thesis an approximate method is developed which predicts the aerodynamic force on a cone-cylinder body in axially symmetric transonic now. The method places more emphasis on the physics of the now than on the mathematical rigors of solving the typical reduced non-linear transonic equation of motion. Under the assumption that the now is that of a steady, irrotational, inviscid, compressible gas, the body pressures are determined and the associated force defined. Recognizing that the transonic pressures are influenced by the character of the subsonic compressible pressures, which are obtained in this analysis through Gothert’s Rule, it is then mandatory that the incompressible case be defined with the best possible accuracy. Comparisons with experiments indicate that the classical method (axially distributed sources and sinks) does not provide this required accuracy. Thus the surface distributed vortex ring theory is used in the present analysis to obtain the incompressible body pressures. Gothert’s Rule, which represents a linear solution for the subsonic case, is known to be applicable up to a limit value of tree stream Mach number. An investigation is carried out herein to determine both the correct form of the rule and its limits of applicability. As a result of this investigation, it is concluded that the upper limit is the lower free stream critical Mach number. Also, at this Mach number, a solution is immediately available tor the lower limit of the transonic range of Mach number. In solving the transonic problem the law or stationarity of local Mach number is of fundamental importance. For an assumed isentropic flow over the body, and for sonic conditions being present at some point on the surface, the body pressures can be described in the ratio p_L/p*. Here p_L is the local surface pressure and p* is the sonic (body) pressure. Through the stationarity law, this ratio is recognized as an invariant for transonic speeds so long as the flow field remains essentially irrotational. Thus any change in local pressure is only a function of the free stream Mach number for any given body position. By this approach, the pressure distribution is defined for a range of Mach number from below to above the sonic stream value. The method is then capable of prediction for almost all of the transonic range of Mach number. It is only when the head shock baa significant curvature, causing the now adjacent to the body to be rotational, that the method fails. Though the procedure developed here is not capable of spanning the entire transonic range, it does provide a wider range of applicability than other known theories. Finally, for this problem, a correlation of transonic pressure drag data is formulated. This correlation is founded on physical interpretation and is not limited to the usual transonic similarity restrictions. In fact, to the author's knowledge, this is the first known such correlation tor axially symmetric flow covering the range of body sizes and Mach numbers considered in this investigation. In so far as is practicable the results obtained in this thesis have been compared to available experimental results. In particular, the drag data from this analysis compare closely with experimental transonic values. Experiment bears out the conclusion that the upper limit for linear theory is the lower critical tree stream Mach number. And, the pressures determined by the vortex ring theory agrees well with the low-speed experimental results obtained by the author. / Ph. D.

The effects of freestream turbulence on the total pressure and total temperature in the wake of a cooled transonic turbine cascade with heated flow are presented in this thesis. The experiment was conducted in the Virginia Tech Cascade Wind Tunnel. A dual hot wire aspirating probe was used to make high frequency, unsteady total pressure and temperature measurements. The probe design was modified to be used in a high temperature environment. The flow was heated to temperatures exceeding 140°C and the turbine blades were actively cooled using gaseous nitrogen to maintain a gas to blade temperature ratio between 1.3 and 1.4. A turbulence screen was used to change the freestream turbulence from 3.3% to 7.5%. Mean and turbulent total pressure and temperature quantities are presented. The higher freestream turbulence resulted in lower total pressure and total temperature turbulence intensities in the wakes of the turbine blades. The freestream turbulence level had no measurable effect on the blade losses. / Master of Science

The aerodynamic and thermodynamic performance of an aircraft gas turbine directly affects the fuel consumption of the engine and the life of the turbine components. Hence, it is important to be able to understand and predict the fluid flow and heat transfer in turbine blades to enable the modifications and improvements in the design process. The use of numerical experiments for the above purposes is becoming increasingly common. The present thesis is involved with the development of a flow solver for turbine flow and heat transfer computations. A 3-D Navier-Stokes code, the Moore Elliptic Flow Program (MEFP) is used to calculate steady flow and heat transfer in turbine rotor cascades. Successful calculations were performed on two different rotor profiles using a one-equation q-L transitional turbulence model. A series of programs was developed for the post-processing of the output from the flow solver. The calculations revealed details of the flow including boundary layer development, trailing edge shocks, flow transition and stagnation and peak heat transfer rates. The calculated pressure distributions, losses, transition ranges, boundary layer parameters and peak heat transfer rates to the blade are compared with the available experimental data. The comparisons indicate that the q-L transitional turbulence model is successful in predicting flows in transonic turbine blade rows. The results also indicate that the calculated loss levels are independent of the gridding used while the heat transfer rate predictions improve with finer grids. / Master of Science

A comprehensive study of the flowfield through a two-dimensional cascade of the high pressure turbine blades of a jet engine is presented. The main interest is the measurement and prediction of the mass-averaged total pressure losses. Other experiments, such as flow visualization, are aimed at the validation of the code that was used to obtain the numerical results and also to further knowledge about the details of the loss generation. The experimental studies were carried out on a cascade of eleven blades in a blow-down tunnel. Total pressure measurements were taken upstream of the cascade and also by traversing on downstream planes. The static pressures needed for the mass averaging and the probe bow shock correction were obtained by pressure taps on the cascade tunnel side wall. The static pressure was also measured on the surface of some instrumented blades. Shadowgraph pictures were taken for study of the trailing edge shock structure and for the turbulent transition location. A single-plate interferometer technique was used for density field measurements. The major goal of the numerical studies was the prediction of the mass-averaged total pressure losses, but all other measured quantities were also generated from the computed flowfield. A critical issue was the generation of a proper grid. For the studied type of flow, a non-periodic C-type grid turned out to be the most advantageous. For use in the moderately compressible attached turbulent boundary layer, a Clauser-type eddy viscosity model was developed and tested. In the trailing edge and wake region, the Baldwin-Lomax model was used. Good agreement of calculations and measurements was obtained for the blade surface and cascade tunnel side wall static pressures, the trailing edge shock structure, and the density field. The agreement between the measured and calculated total pressure drop profiles was not quite as good; however, that quantity is known to be difficult to predict accurately. The mass-averaged total pressure loss coefficient, calculated from the total pressure drop profiles, was again in good agreement with the measurements. The difference between the measured and computed total pressure drop profiles suggested that the Baldwin-Lomax model underpredicted the eddy viscosity in the trailing edge region. / Ph. D.