Thesis (M.S.)--University of Wisconsin--Madison, 1966. / eContent provider-neutral record in process. Description based on print version record. Bibliography: l. 126-128.

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

An existing Navier-Stokes code (MEFP) was used to calculate developing flow and heat transfer in turbine tip gaps. Successful calculations of the heat transfer to a model turbine blade tip were obtained with a Prandtl mixing length turbulence model. The calculations revealed details of the flow development including recirculation and reattachment on the blade tip surface. The calculated heat transfer distributions were in good agreement with experimental data. A combined solution of the energy equation in the tip gap flow and in the rotor blade tip gave tip temperature distributions. An independent computational study, using the same numerics as MEFP but a separate new computer program, was also performed to investigate the numerical accuracy of heat transfer calculations for fully developed flow. A literature survey of gas turbine blade materials and factors influencing turbine tip blade life was performed. Approximate temperature ranges for the significant blade life reduction mechanisms, hot corrosion, oxidation, and melting were determined. calculations for typical jet engine conditions, In the present a maximum tip temperature of 1488 K was predicted which would lead to high oxidation rates for present day turbine blade alloys. / M.S.

A panel method derived from inviscid irrotational flow theory and utilizing hyperboloid panels is developed and applied to the simulation of steady fully wetted flows inside water-jet pumps and rocket engine inducers. The source and dipole influence coefficients of the hyperboloid panels are computed using Gauss quadrature. The present method solves the boundary value problem subject to a uniform inflow directly by discretizing the blade, casing/shroud and hub geometries with panels. The Green's integral equation and the influence coefficients for the water-jet/inducer problem are defined and solved by allocating constant strength sources and dipoles on the blade, hub and casing surfaces and constant strength dipoles on the shed wake sheets from the rotor/ stator blades. The rotor- stator interaction is accomplished using an iterative procedure which considers the effects between the rotor and the stator, via circumferentially averaged induced velocities. Finally, the hydrodynamic performance predictions for the water-jet pump and the inducer from the present method are validated against existing experimental data and numerical results from Reynolds Averaged Navier- Stokes (RANS) solvers. / text

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This research report documents the various types of Data Acquisition Systems in use for testing jet aircraft engines. The cost trade offs and design considerations are explored for systems which employ a digital computer as the prime recording/processing element. The digital computer has revolutionized the data acquisition field, particularly in the testing of high performance jet engines. Test data can be acquired, processed, converted to engineering units, and out via high speed line printers and cathode ray tubes (CRT's). The data acquisition system operates on-line, and interleaves the random requests for data from multiple test cells by using a specially designed software system and multi-processing capability of the high speed digital computer. All test data must be traceable to The National Bureau of Standards, which required that all calibration standards also be traceable. Primary and secondary calibration methods are discussed and examples of the mathematical processes for conversion of the raw data to meaningful results are presented. Data Acquisition Systems for jet engine testing can be logically grouped into two main categories, with the determining factor being the type of test to be conducted. Production engine testing requires rapid setup, calibration, and fast data turn around, particularly for modern automated test facilities. Development engine testing requires a large number of data channels, infrequent setup, and complete software for extensive engine performance calculations. Both types of Data Acquisition Systems have been designed and built by Pratt & Whitney Aircraft and are used as examples of the techniques described in this report.

This study examines NACA's organization in the light of Alfred D. Chandler's <i>Strategy and Structure</i>. It analyzes the agency's administration. NACA's strategy of maximizing existing technology and its committee's structure were the key elements in its failure to develop jet propulsion in the early 1940s. We will focus first on NACA and its organization. The second chapter will describe jet propulsion, particularly the acquisition of a Whittle engine from England and General Arnold 's role in keeping NACA out of the development of the Whittle engine in the United States. The third chapter will concentrate on the reasons that combined and led to the difficulties of NACA in the mid-forties and the 1950s. That chapter will look at the rise of the aviation industry, the criticism it expressed against NACA, and finally NACA's strategy as one of the causes of failure. / Master of Arts

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