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Performance evaluation of reduced-chord rotor blading as applied to J73 two-stage turbine 5: effect of inlet pressure on over-all performance at design speed and inlet temperature of 700 degrees R

Description: Report presenting an investigation of a reduced-chord multistage turbine at design speed and various turbine-inlet pressures from 12 to 40 inches of mercury absolute. At each inlet pressure, the turbine was operated over a range of overall turbine total-pressure ratios; turbine-inlet temperature was maintained at 700 degrees R. The results indicated that no appreciable effect on turbine overall performance was observed over the range of turbine-inlet total pressures investigated.
Date: July 5, 1957
Creator: Schum, Harold J.
Partner: UNT Libraries Government Documents Department

Design and Experimental Investigation of Light-Weight Bases for Air-Cooled Turbine Rotor Blades

Description: Memorandum presenting an investigation of the problem of air-cooled turbine rotor-blade weight reduction for aircraft gas turbines as part of the NACA turbine-cooling research program. A bulb-root-type blade-base design was achieved which utilized forming and brazing techniques in its manufacture and was combined with an axial-flow turbojet engine aerodynamic profile. Results regarding the sheet-metal bulb-root-type air-cooled base designs, air-cooled pinned-type base design, and some weight reduction considerations are provided.
Date: July 2, 1954
Creator: Freche, John C. & McKinnon, Roy A.
Partner: UNT Libraries Government Documents Department

The Efficiency of Combustion Turbines With Constant-Pressure Combustion

Description: "Of the two fundamental cycles employed in combustion turbines, namely, the explosion (or constant-volume) cycle and the constant-pressure cycle, the latter is considered more in detail and its efficiency is derived with the aid of the cycle diagrams for the several cases with adiabatic and isothermal compression and expansion strokes and with and without utilization of the exhaust heat. Account is also taken of the separate efficiencies of the turbine and compressor and of the pressure losses and heat transfer in the piping. The results show that without the utilization of the exhaust heat the efficiencies for the two cases of adiabatic and isothermal compression is offset by the increase in the heat supplied" (p. 1).
Date: April 1941
Creator: Piening, Werner
Partner: UNT Libraries Government Documents Department

Experimental Investigation of a 7-Inch-Tip-Diameter Transonic Turbine

Description: Memorandum presenting the overall performance results obtained for a 7-inch transonic turbine and a comparison with the results of a 14-inch turbine of geometrically similar design that had been previously investigated. The peak efficiency obtained with the turbine was 0.85, which was 2 points lower than that obtained previously with the 14-inch turbine.
Date: January 16, 1958
Creator: Whitney, Warren J. & Wintucky, William T.
Partner: UNT Libraries Government Documents Department

Design and Experimental Investigation of a Single-Stage Turbine With a Rotor Entering Relative Mach Number of 2

Description: Memorandum presenting a design and experimental investigation of a single-stage supersonic turbine. The turbine was designed for a rotor entering relative Mach number of 2. Results regarding the overall turbine performance, outer-wall static-pressure variation, momentum-loss considerations, effect of rotor modifications, and supersonic starting are provided.
Date: September 15, 1958
Creator: Moffitt, Thomas P.
Partner: UNT Libraries Government Documents Department

Investigation of turbines suitable for use in a turbojet engine with high compressor pressure ratio and low compressor-tip speed 6: experimental performance of two-stage turbine

Description: The brake internal efficiency of the highly loaded two-stage turbine was 0.79 equivalent design work and speed. The maximum brake internal efficiency was 0.84. A radial survey revealed these major defects: (1) the first-rotor throat area was too large, and a large area of underturned flow existed near the tip;(2) considerable underturning existed at the second-stator outlet; and (3) tangential components of velocity at the turbine outlet amounted to 2.5 points in turbine efficiency.
Date: August 20, 1956
Creator: Davison, Elmer H.; Schum, Harold J. & Petrash, Donald A.
Partner: UNT Libraries Government Documents Department

Mechanical design analysis of several noncritical air-cooled turbine disks and a corrugated-insert air-cooled turbine rotor blade

Description: Report presenting a series of turbine wheel designs made for a turbojet engine with an axial-flow compressor. The designs were made to determine the influence of air-cooling on overall engine design, weight, critical-material content, and complexity and to assist with construction of research model full-scale turbines.
Date: July 22, 1953
Creator: Moseson, Merland L.; Krasner, Morton H. & Ziemer, Robert R.
Partner: UNT Libraries Government Documents Department

One-Dimensional Analysis of Choked-Flow Turbines

Description: "Turbines for most applications requiring high work output per stage have one or more blade rows which are choked. This analysis indicated that the area ratios and equivalent blade speed are the controlling factors in the design and operation of such turbines. Six criteria are stated that will aid in establishing from test data of multistage turbines which blade rows are choked and which are not" (p. 1).
Date: 1953
Creator: English, Robert E. & Cavicchi, Richard H.
Partner: UNT Libraries Government Documents Department

Performance Evaluation of Reduced-Chord Rotor Blading as Applied to J73 Two-Stage Turbine 6: Stage Performance With Standard Rotor Blading at Inlet Conditions of 35 Inches of Mercury Absolute and 700 Degrees R

Description: Report presenting an evaluation of the stage performance of the J73 turbine on the basis of previous performance investigations of the first stages alone and of the two-stage turbines with standard and with reduced-chord rotor blading for a range of rotational speeds and pressure ratios. The results demonstrated that both of the two-stage turbines and first-stage turbines had comparable performance.
Date: July 11, 1957
Creator: Davison, Elmer H. & Schum, Harold J.
Partner: UNT Libraries Government Documents Department

Performance of 19XB-2A Gas Turbine 1 - Effect of Pressure Ratio and Inlet Pressure on Turbine Performance for an Inlet Temperature of 800 Degrees R

Description: "An investigation of the 19XB-2A gas turbine is being conducted at the Cleveland laboratory to determine the effect on turbine performance of various inlet pressures, inlet temperatures, pressure ratios, and wheel speeds. The engine of which this turbine is a component is designed to operate at an air flow of 30 pounds per second at a compressor rotor speed of 17,000 rpm at sea-level conditions. At these conditions the total-pressure ratio is 2.08 across the turbine and the turbine inlet total temperature is 2000 degrees R" (p. 1).
Date: December 26, 1946
Creator: Kohl, Robert C. & Larkin, Robert G.
Partner: UNT Libraries Government Documents Department

Performance Evaluation of Reduced-Chord Rotor Blading as Applied to J73 Two-Stage Turbine

Description: "The multistage turbine from the J73 turbojet engine has previously been investigated with standard and with reduced-chord rotor blading in order to determine the individual performance characteristics of each configuration over a range of over-all pressure ratio and speed. Because both turbine configurations exhibited peak efficiencies of over 90 percent, and because both units had relatively wide efficient operating ranges, it was considered of interest to determine the performance of the first stage of the turbine as a separate component. Accordingly, the standard-bladed multistage turbine was modified by removing the second-stage rotor disk and stator and altering the flow passage so that the first stage of the unit could be operated independently" (p. 1).
Date: July 11, 1957
Creator: Schurn, Harold J.
Partner: UNT Libraries Government Documents Department

Investigation of a Transonic Turbine Designed for a Maximum Rotor-Blade Suction-Surface Relative Mach Number of 1.57

Description: "A transonic turbine designed for a maximum blade-surface relative Mach number of 1.57 was investigated experimentally. The performance of the turbine is compared with that of three other transonic turbines that were previously investigated" (p. 1).
Date: October 11, 1954
Creator: Whitney, Warren J.; Wong, Robert Y. & Monroe, Daniel E.
Partner: UNT Libraries Government Documents Department

Offshore Wind Turbines - Estimated Noise from Offshore Wind Turbine, Monhegan Island, Maine: Environmental Effects of Offshore Wind Energy Development

Description: Deep C Wind, a consortium headed by the University of Maine will test the first U.S. offshore wind platforms in 2012. In advance of final siting and permitting of the test turbines off Monhegan Island, residents of the island off Maine require reassurance that the noise levels from the test turbines will not disturb them. Pacific Northwest National Laboratory, at the request of the University of Maine, and with the support of the U.S. Department of Energy Wind Program, modeled the acoustic output of the planned test turbines.
Date: November 23, 2010
Creator: Aker, Pamela M.; Jones, Anthony M. & Copping, Andrea E.
Partner: UNT Libraries Government Documents Department

Cast Alloys for Advanced Ultra Supercritical Steam Turbines

Description: The proposed steam inlet temperature in the Advanced Ultra Supercritical (A-USC) steam turbine is high enough (760 °C) that traditional turbine casing and valve body materials such as ferritic/martensitic steels will not suffice due to temperature limitations of this class of materials. Cast versions of several traditionally wrought Ni-based superalloys were evaluated for use as casing or valve components for the next generation of industrial steam turbines. The full size castings are substantial: 2-5,000 kg each half and on the order of 100 cm thick. Experimental castings were quite a bit smaller, but section size was retained and cooling rate controlled to produce equivalent microstructures. A multi-step homogenization heat treatment was developed to better deploy the alloy constituents. The most successful of these cast alloys in terms of creep strength (Haynes 263, Haynes 282, and Nimonic 105) were subsequently evaluated by characterizing their microstructure as well as their steam oxidation resistance (at 760 and 800 °C).
Date: May 1, 2010
Creator: Holcomb, G. R.; Wang, P.; Jablonski, P. D. & Hawk, J. A.
Partner: UNT Libraries Government Documents Department

Cooling of Gas Turbines, 2, Effectiveness of Rim Cooling of Blades

Description: An analysis is presented of rim cooling of gas-turbine blades; that is, reducing the temperature at the base of the blade (wheel rim), which cools the blade by conduction alone. Formulas for temperature and stress distributions along the blade are derived and, by the use of experimental stress-rupture data for a typical blade alloy, a relation is established between blade life (time for rupture), operating speed, and amount of rim cooling for several gas temperatures. The effect of blade parameter combining the effects of blade dimensions, blade thermal conductivity, and heat-transfer coefficient is determined. The effect of radiation on the results is approximated. The gas temperatures ranged from 1300F to 1900F and the rim temperature, from 0F to 1000F below the gas temperature. This report is concerned only with blades of uniform cross section, but the conclusions drawn are generally applicable to most modern turbine blades. For a typical rim-cooled blade, gas temperature increases are limited to about 200F for 500F of cooling of the blade base below gas temperature, and additional cooling brings progressively smaller increases. In order to obtain large increases in thermal conductivity or very large decreases in heat-transfer coefficient or blade length or necessary. The increases in gas temperature allowable with rim cooling are particularly small for turbines of large dimensions and high specific mass flows. For a given effective gas temperature, substantial increases in blade life, however, are possible with relatively small amounts of rim cooling.
Date: March 18, 1947
Creator: Wolfenstein, Lincoln; Meyer, Gene L. & McCarthy, John S.
Partner: UNT Libraries Government Documents Department

Investigation of a high-temperature single-stage turbine suitable for air cooling and turbine stator adjustment 2: performance of vortex turbine at various stator settings

Description: Report presenting a consideration of a mode of engine operation that requires operational flexibility of the turbine, which requires that the turbine stator and exhaust nozzle area are adjusted to maintain a fixed compressor operating point. Results regarding turbine performance and comparison of predicted and experimental turbine performance are provided.
Date: August 3, 1954
Creator: Heaton, Thomas R.; Holeski, Donald E. & Forrette, Robert E.
Partner: UNT Libraries Government Documents Department

Experimental Permeability Measurements on a Strut-Supported Transpiration-Cooled Turbine Blade with Stainless-Steel Shell made by the Federal-Mogul Corporation under Bureau of Aeronautics Contract N0as 51613-C

Description: A turbine blade with a porous stainless-steel shell sintered to a supporting steel strut has been fabricated for tests at the NACA by Federal-Mogul Corporation under contract from the Bureau of Aeronautics, Department of the Navy. The apparent permeability of this blade, on the average, more nearly approaches the values specified by the NAGA than did two strut-supported bronze blades in a previous investigation. Random variations of permeability in the present blade are substantialy greater than those of the bronze blades, but projected improvements in certain phases of the fabrication process are expected to reduce these variations.
Date: April 30, 1954
Creator: Richards, Hadley T.
Partner: UNT Libraries Government Documents Department

Turbine performance characteristics of a python turbine-propeller engine investigated in altitude wind tunnel

Description: Report presenting the performance of the turbine component of a Python turbine-propeller engine with four tail-pipe configurations determined over a range of altitudes, engine speeds, and fuel flows. Results regarding the corrected turbine speed, corrected turbine enthalpy drop, pressure ratio, and efficiency were established on a turbine-characteristic plot.
Date: May 1, 1951
Creator: Farley, John M. & Prince, William R.
Partner: UNT Libraries Government Documents Department

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Nozzles and Three Rotor-Blade Designs

Description: From Summary: "A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzles and three rotor-blade designs. The performance was evaluated in terms of brake, rotor, and blade efficiencies at pressure ratios of 8, 15 (design), and 20. The blade efficiencies with the two nozzles are compared with those obtained with four other nozzles previously investigated with the same three rotor-blade designs."
Date: September 9, 1949
Creator: Schum, Harold J. & Whitney, Warren J.
Partner: UNT Libraries Government Documents Department

Use of effective momentum thickness in describing turbine rotor-blade losses

Description: Report presenting a discussion of the use of an effective rotor-blade momentum thickness in describing rotor-blade loss characteristics. A derivation of the necessary equations is presented for obtaining momentum thickness for given overall turbine performance, stator performance, and rotor geometric quantities.
Date: May 14, 1956
Creator: Stewart, Warner L.; Whitney, Warren J. & Miser, James W.
Partner: UNT Libraries Government Documents Department

Determination of elastic stresses in gas-turbine disks

Description: A method is presented for the calculation of elastic stresses in symmetrical disks typical of those of a high-temperature gas turbine. The method is essentially a finite-difference solution of the equilibrium and compatibility equations for elastic stresses in a symmetrical disk. Account can be taken of point-to-point variations in disk thickness, in temperature, in elastic modulus, in coefficient of thermal expansion, in material density, and in Poisson's ratio. No numerical integration or trial-and-error procedures are involved and the computations can be performed in rapid and routine fashion by nontechnical computers with little engineering supervision. Checks on problems for which exact mathematical solutions are known indicate that the method yields results of high accuracy. Illustrative examples are presented to show the manner of treating solid disks, disks with central holes, and disks constructed either of a single material or two or more welded materials. The effect of shrink fitting is taken into account by a very simple device.
Date: 1947
Creator: Manson, S. S.
Partner: UNT Libraries Government Documents Department

Vibrational modes of several hollow turbine blades and of solid turbine blade of similar aerodynamic design

Description: Experimental study to determine the vibrational modes of several hollow turbine blades and a solid turbine blade of similar aerodynamic design. Results regarding the significance of nodal patterns, vibrational modes of six blade types, and the probability of excitation in some of the blades are provided.
Date: October 3, 1949
Creator: Kemp, R. H. & Shifman, J.
Partner: UNT Libraries Government Documents Department