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Brown, W Byron

Publications and source records attributed to Brown, W Byron.

18 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1953

Comparison of Theoretically and Experimentally Determined Effects of Oxide Coatings Supplied by Fuel Additives on Uncooled Turbine-blade Temperature During Transient Turbojet-engine Operation

An analysis was made to permit the calculation of the effectiveness of oxide coatings in retarding the transient heat flow into turbine blades when the combustion gas temperature of a turbojet engine is suddenly changed. The analysis is checked with experimental data obtained from a turbojet engine whose blades were coated with two different coating materials (silicon dioxide and boric oxide) by adding silicone oil and tributyl borate to the engine fuel. The very thin coatings (approximately 0.001 in.) that formed on the blades produced a negligible effect on the turbine-blade transient temperature response. With the analysis discussed here, it was possible to predict the turbine rotor-blade temperature response with a maximum error of 40 F.

PROTECTIVE COATINGS↗
NASA NTRS1952

Solutions of laminar-boundary-equations which result in specific-weight-flow profiles locally exceeding free-stream values

Revised solutions of the laminar-boundary-layer equations for cases which involved cooling at the wall combined with large pressure gradients in the main stream produced specific-weight-flow profiles which locally exceeded free-stream values. Heat-transfer and friction coefficients, boundary-layer thicknesses, and velocity, temperature, and specific-weight-flow distributions resulting from the revised solutions are presented for Euler numbers of 0.5 and 1, stream-to-wall temperature ratios of 2 and 4, and cooling-air flow rates through porous walls designated by flow parameters of 0, -0.5, and -1.

HEAT TRANSFER, AERODYNAMIC↗
NASA NTRS1952

Analysis of temperature distribution in liquid-cooled turbine blades

The temperature distribution in liquid-cooled turbine blades determines the amount of cooling required to reduce the blade temperature to permissible values at specified locations. This report presents analytical methods for computing temperature distributions in liquid-cooled turbine blades, or in simplified shapes used to approximate sections of the blade. The individual analyses are first presented in terms of their mathematical development. By means of numerical examples, comparisons are made between simplified and more complete solutions and the effects of several variables are examined. Nondimensional charts to simplify some temperature-distribution calculations are also given.

Livingood, John N B↗
NASA NTRS1950

Extension of boundary-layer heat-transfer theory to cooled turbine blades

An equation for average heat transfer of a surface was derived when the boundary layer changed from laminar to turbulent. Influences on the heat transfer through a laminar boundary layer of Mach number, temperature ratio (gas temperature divided by wall temperature), and exponents of gas-property temperature relations were shown to be relatively small for air with Mach numbers less than 2 and temperature ratios between 1 and 4. Good agreement was obtained with experimental results from cylinders, an airfoil, and turbine blades.

HEAT TRANSFER, AERODYNAMIC↗
NASA NTRS1950

Analytical determination of local surface heat-transfer coefficients for cooled turbine blades from measured metal temperatures

Analytical methods are presented for the determination of local values of outside and inside heat-transfer coefficients and effective gas temperatures by use of turbine-blade-temperature measurements. The methods are derived for a number of configurations that can be applied to typical cooled-turbine-blade shapes as well as to other types of heat-transfer apparatus.

TURBINE COOLING↗
NASA NTRS1950

Analysis of spanwise temperature distribution in three types of air-cooled turbine blade

Methods for computing spanwise blade-temperature distributions are derived for air-cooled hollow blades, air-cooled hollow blades with inserts, and air-cooled blades containing internal cooling fins. Individual and combined effects on spanwise blade-temperature distributions of cooling-air and radial heat conduction are determined. In general, the effects of radiation and radial heat conduction were found to be small and the omission of these variations permitted the construction of nondimensional charts for use in determining spanwise temperature distribution through air-cooled turbine blades. An approximate method for determining the allowable stress-limited blade-temperature distribution is included, with brief accounts of a method for determining the maximum allowable effective gas temperatures and the cooling-air requirements. Numerical examples that illustrate the use of the various temperature-distribution equations and of the nondimensional charts are also included.

Livingood, John N B↗
NASA NTRS1949

Design and performance of family of diffusing scrolls with mixed-flow impeller and vaneless diffuser

A family of diffusing scrolls was designed for use with a mixed-flow impeller and a small-diameter vaneless diffuser. The design theory, intended to maintain a uniform pressure around the scroll inlet, permits determination of the position of scroll cross sections of preassigned area by considering the radial variation in fluid density and the effects of friction along the scroll. Inasmuch as the design method leaves the cross-sectional shape undetermined, the effect of certain variations in scroll shape was investigated by studying scrolls having angles of divergence (of the scroll walls downstream of the entrance section) of 24 degrees, 40 degrees, and 80 degrees. A second 80 degree scroll was of asymmetrical construction and a third was plaster-cast instead of sand-cast. Each scroll was tested as a compressor component at actual impeller tip speeds of 700 to 1300 feet per second from full throttle to surge.

Brown, W Byron↗
NASA NTRS1948

Cooling of gas turbines IX : cooling effects from use of ceramic coatings on water-cooled turbine blades

The hottest part of a turbine blade is likely to be the trailing portion. When the blades are cooled and when water is used as the coolant, the cooling passages are placed as close as possible to the trailing edge in order to cool this portion. In some cases, however, the trailing portion of the blade is so narrow, for aerodynamic reasons, that water passages cannot be located very near the trailing edge. Because ceramic coatings offer the possibility of protection for the trailing part of such narrow blades, a theoretical study has been made of the cooling effect of a ceramic coating on: (1) the blade-metal temperature when the gas temperature is unchanged, and (2) the gas temperature when the metal temperature is unchanged. Comparison is also made between the changes in the blade or gas temperatures produced by ceramic coatings and the changes produced by moving the cooling passages nearer the trailing edge. This comparison was made to provide a standard for evaluating the gains obtainable with ceramic coatings as compared to those obtainable by constructing the turbine blade in such a manner that water passages could be located very near the trailing edge.

HEAT TRANSMISSION - METALS↗