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  Partner: UNT Libraries Government Documents Department
 Serial/Series Title: NACA Research Memorandums
 Collection: Technical Report Archive and Image Library
Effect of fuels on screaming in 200-pound-thrust liquid-oxygen - fuel rocket engine

Effect of fuels on screaming in 200-pound-thrust liquid-oxygen - fuel rocket engine

Date: June 22, 1956
Creator: Pass, Isaac & Tischler, Adelbert O
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Effect of fuselage circumferential inlet location on diffuser-discharge total-pressure profiles at supersonic speeds

Effect of fuselage circumferential inlet location on diffuser-discharge total-pressure profiles at supersonic speeds

Date: October 29, 1956
Creator: Kremzier, Emil J & Wasserbauer, Joseph F
Description: An experimental investigation of the effect of angle of attack and inlet corrected air flow on diffuser-discharge total-pressure profiles of inlets located in various circumferential positions on a fuselage was conducted at supersonic speeds. Results indicated that the diffuser total-pressure profiles for a bottom inlet were least affected by angle of attack on distortion level was obtained with a side inlet. Variation in distortion for top inlets with angle of attack was confined to the supercritical range of inlet operation.
Contributing Partner: UNT Libraries Government Documents Department
Effect of geometry on secondary flows in blade rows

Effect of geometry on secondary flows in blade rows

Date: October 16, 1952
Creator: Hansen, A G; Costello, G R & Herzig, H Z
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Effect of fuselage fences on the angle-of-attack supersonic performance of a top-inlet - fuselage

Effect of fuselage fences on the angle-of-attack supersonic performance of a top-inlet - fuselage

Date: January 19, 1955
Creator: Kremzier, Emil J & Campbell, Robert C
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Characteristics of a canard-type missile configuration with an underslung scoop inlet at Mach numbers from 1.5 to 2.0

Characteristics of a canard-type missile configuration with an underslung scoop inlet at Mach numbers from 1.5 to 2.0

Date: January 30, 1953
Creator: Fradenburgh, Evan A & Campbell, Robert C
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Characteristics of a Hot Jet Discharged from a Jet-Propulsion Engine

Characteristics of a Hot Jet Discharged from a Jet-Propulsion Engine

Date: December 27, 1946
Creator: Fleming, William A.
Description: An investigation of a heated jet was conducted in conjunction with tests of an axial-flow jet-propulsion engine in the Cleveland altitude wind tunnel. Pressure and temperature surveys were made across the jet 10 and 15 feet behind the jet-nozzle outlet of the engine. Surveys were obtained at pressure altitudes of 10,000, 20,000, 30,000, and 40,000 feet with test-section velocities from 30 to 110 feet per second and test-section temperatures from 60 F to -50 F. From measurements taken throughout the operable range of engine speeds, tail-pipe outlet temperatures from 500 F to 1250 F and jet velocities from 400 to 2200 feet per second were obtained. The jet-survey data presented extend the work previously done with low-velocity and low-temperature jets to the region of high velocities and high temperatures. The results obtained agree with previously determined experimental data and with predicted theoretical expressions for the dimensionless transverse velocity and temperature profiles across a jet. The spread of both the temperature and the velocity profiles was very nearly linear. Dimensionless plots of temperature and velocity along the axis of a heated jet agree with experimental results of tests with a cold jet.
Contributing Partner: UNT Libraries Government Documents Department
Characteristics of a hydraulic control determined from transient data obtained with a turbojet engine at altitude

Characteristics of a hydraulic control determined from transient data obtained with a turbojet engine at altitude

Date: June 16, 1954
Creator: Vasu, George; Hinde, William L & Craig, R T
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Characteristics of a wedge with various holder configurations for static-pressure measurements in subsonic gas streams

Characteristics of a wedge with various holder configurations for static-pressure measurements in subsonic gas streams

Date: September 5, 1951
Creator: Gettelman, Clarence C & Krause, Lloyd N
Description: The characteristics of a wedge static-pressure sensing element with various holder configurations were determined and compared with the characteristics of the conventional tube. The probes were tested over a range of Mach number from 0.3 to 0.95 and at various pitch and yaw angles. The investigation showed that the spike-mounted wedge sensing element has a pressure coefficient comparable with the conventional subsonic static-pressure probe and the pressure coefficient of the wedge varied less than that of the conventional probe for corresponding change of yaw angle.
Contributing Partner: UNT Libraries Government Documents Department
Characteristics of flow about axially symmetric isentropic spikes for nose inlets at Mach number 3.85

Characteristics of flow about axially symmetric isentropic spikes for nose inlets at Mach number 3.85

Date: August 19, 1954
Creator: Connors, James F & Wollett, Richard R
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Cooling of Gas Turbines I - Effects of Addition of Fins to Blade Tips and Rotor, Admission of Cooling Air Through Part of Nozzles, and Change in Thermal Conductivity of Turbine Components

Cooling of Gas Turbines I - Effects of Addition of Fins to Blade Tips and Rotor, Admission of Cooling Air Through Part of Nozzles, and Change in Thermal Conductivity of Turbine Components

Date: February 11, 1947
Creator: Brown, Byron
Description: An analysis was developed for calculating the radial temperature distribution in a gas turbine with only the temperatures of the gas and the cooling air and the surface heat-transfer coefficient known. This analysis was applied to determine the temperatures of a complete wheel of a conventional single-stage impulse exhaust-gas turbine. The temperatures were first calculated for the case of the turbine operating at design conditions of speed, gas flow, etc. and with only the customary cooling arising from exposure of the outer blade flange and one face of the rotor to the air. Calculations were next made for the case of fins applied to the outer blade flange and the rotor. Finally the effects of using part of the nozzles (from 0 to 40 percent) for supplying cooling air and the effects of varying the metal thermal conductivity from 12 to 260 Btu per hour per foot per degree Farenheit on the wheel temperatures were determined. The gas temperatures at the nozzle box used in the calculations ranged from 1600F to 2000F. The results showed that if more than a few hundred degrees of cooling of turbine blades are required other means than indirect cooling with fins on the rotor ...
Contributing Partner: UNT Libraries Government Documents Department
Cooling of Gas Turbines, 6, Computed Temperature Distribution Through Cross Section of Water-Cooled Turbine Blade

Cooling of Gas Turbines, 6, Computed Temperature Distribution Through Cross Section of Water-Cooled Turbine Blade

Date: May 1, 1947
Creator: Livingood, John N. B. & Sams, Eldon W.
Description: A theoretical analysis of the cross-sectional temperature distribution of a water-cooled turbine blade was made using the relaxation method to solve the differential equation derived from the analysis. The analysis was applied to specific turbine blade and the studies icluded investigations of the accuracy of simple methods to determine the temperature distribution along the mean line of the rear part of the blade, of the possible effect of varying the perimetric distribution of the hot gas-to -metal heat transfer coefficient, and of the effect of changing the thermal conductivity of the blade metal for a constant cross sectional area blade with two quarter inch diameter coolant passages.
Contributing Partner: UNT Libraries Government Documents Department
Cooling of Gas Turbines, IV - Calculated Temperature Distribution in the Trailing Part of a Turbine Blade Using Direct Liquid Cooling

Cooling of Gas Turbines, IV - Calculated Temperature Distribution in the Trailing Part of a Turbine Blade Using Direct Liquid Cooling

Date: April 18, 1947
Creator: Brown, W. Byron & Monroe, William R.
Description: A theoretical analysis of the temperature distribution through the trailing portion of a blade near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the hot spot temperatures at the trailing edge and influence of design variables. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.
Contributing Partner: UNT Libraries Government Documents Department
Cooling of gas turbines IX : cooling effects from use of ceramic coatings on water-cooled turbine blades

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

Date: October 13, 1948
Creator: Brown, W Byron & Livingood, John N B
Description: 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.
Contributing Partner: UNT Libraries Government Documents Department
Cooling of gas-turbines VII : effectiveness of air cooling of hollow turbine blades with inserts

Cooling of gas-turbines VII : effectiveness of air cooling of hollow turbine blades with inserts

Date: October 20, 1947
Creator: Bressman, Joseph R & Livingood, John N B
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Cooling of ram jets and tail-pipe burners : analytical method  for determining temperatures of combustion chamber having annular cooling passage

Cooling of ram jets and tail-pipe burners : analytical method for determining temperatures of combustion chamber having annular cooling passage

Date: March 21, 1950
Creator: Koffel, William K; Stamper, Eugene & Sanders, Newell D
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Cooling performance and structural reliability of a modified corrugated-insert air-cooled turbine blade with an integrally cast shell and base

Cooling performance and structural reliability of a modified corrugated-insert air-cooled turbine blade with an integrally cast shell and base

Date: January 21, 1957
Creator: Freche, John C & Schum, Eugene F
Description: A modified corrugated-insert blade with integrally cast shell and base was developed. This blade was as light as a conventional fabricated corrugated-insert blade. Of four test blades operated in a full-scale turbojet engine, one failed after about 15 hours operation at an inlet gas temperature of 1670 degrees F, a coolant-flow ratio of 0.0064, and a 1/3-span centrifugal stress of approximately 28,000 psi. Three other test blades ran for approximately 16, 31, and 36 hours without failure at similar conditions.
Contributing Partner: UNT Libraries Government Documents Department
Correlation between hydrogen pressure and protective action of additives in the molten sodium hydroxide - nickel system

Correlation between hydrogen pressure and protective action of additives in the molten sodium hydroxide - nickel system

Date: February 9, 1956
Creator: May, Charles E
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Coolant-Flow Calibrations of Three Simulated Porous Gas-Turbine Blades

Coolant-Flow Calibrations of Three Simulated Porous Gas-Turbine Blades

Date: March 8, 1951
Creator: Esger, Jack B. & Lea, Alfred L.
Description: An investigation was conducted at the NACA Lewis laboratory to determine whether simulated porous gas-turbine blades fabricated by the Eaton Manufacturing Company of Cleveland, Ohio would be satisfactory with respect to coolant flow for application in gas-turbine engines. These blades simulated porous turbine blades by forcing the cooling air onto the blade surface through a large number of chordwise openings or slits between laminations of sheet metal or wire. This type of surface has a finite number of openings, whereas a porous surface has an almost infinite number of smaller openings for the coolant flow. The investigation showed that a blade made of sheet-metal laminations stacked on a support member that passed up through the coolant passage was completely unsatisfactory because of extremely poor coolant flow distribution over the blade surface. The flow distribution for two wire-wound blades was more uniform, but the pressure drop between the coolant supply pressure and the local pressure on the outside of the blades was too low by a factor ranging from 3 to 3.5 for the required coolant flow rates. The pressure drop could be increased by forcing the wires closer together during blade fabrication.
Contributing Partner: UNT Libraries Government Documents Department
Cooling characteristics of an experimental tail-pipe burner with an annular cooling-air passage

Cooling characteristics of an experimental tail-pipe burner with an annular cooling-air passage

Date: February 26, 1952
Creator: Kaufman, Harold R & Koffel, William K
Description: The effects of tail-pipe fuel-air ratio (exhaust-gas temperatures from approximately 3060 degrees to 3825 degrees R), radial distributiion of tail-pipe fuel flow, and mass flow of combustion gas and the inside wall were determined for an experimental tail-pipe burner cooled by air flowing through and insulated cooling-air to combustion gas mass flow from 0.066 to 0.192 were also determined.
Contributing Partner: UNT Libraries Government Documents Department
Considerations in the adaptation of low-cost fuels to gas-turbine-powered commercial aircraft

Considerations in the adaptation of low-cost fuels to gas-turbine-powered commercial aircraft

Date: October 1, 1953
Creator: Barnett, Henry C & Mccafferty, Richard J
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Cooling characteristics of a transpiration-cooled afterburner with a porous wall of brazed and rolled wire cloth

Cooling characteristics of a transpiration-cooled afterburner with a porous wall of brazed and rolled wire cloth

Date: August 19, 1954
Creator: Koffel, William K
Description: None
Contributing Partner: UNT Libraries Government Documents Department
Criterions for prediction and control of ram-jet flow pulsations

Criterions for prediction and control of ram-jet flow pulsations

Date: May 16, 1951
Creator: Sterbentz, William H & Evvard, John C
Description: None
Contributing Partner: UNT Libraries Government Documents Department
The Crystal Structure at Room Temperature of Six Cast Heat-Resisting Alloys

The Crystal Structure at Room Temperature of Six Cast Heat-Resisting Alloys

Date: June 3, 1947
Creator: Rosenbaum, Burt M.
Description: The crystal structures of alloys 61, X-40,X-50, 422-19, 6059, and Vitallium, derived from x-ray diffraction, are discussed. The alloys have been, or are being considered for use in gas turbine applications. The predominant phase was a solid solution of the face centered cubic type of the principal constituent elements.The lattice parameters were found to be between 3.5525 and 3.5662.
Contributing Partner: UNT Libraries Government Documents Department
Approximate relative-total-pressure losses of an infinite cascade of supersonic blades with finite leading-edge thickness

Approximate relative-total-pressure losses of an infinite cascade of supersonic blades with finite leading-edge thickness

Date: March 3, 1950
Creator: Klapproth, John F
Description: None
Contributing Partner: UNT Libraries Government Documents Department