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Aerodynamic study of a wing-fuselage combination employing a wing swept back 63 degrees : characteristics for symmetrical wing sections at high subsonic and moderate supersonic Mach numbers

Description: From Summary: "Results of wind-tunnel tests are presented for a wing with the leading edge swept back 63^o and of symmetrical section in combination with a body at Mach numbers from 0.5 to 0.95 and from 1.09 to 1.51."
Date: July 7, 1949
Creator: Mas, Newton A
Partner: UNT Libraries Government Documents Department

Air-Stream Surveys in the Vicinity of the Tail of a 1/8.33-Scale Powered Model of the Republic XF-12 Airplane

Description: The XF-12 airplane was designed by Republic Aviation Corporation to provide the Army Air Forces with a high performance, photo reconnaissance aircraft. A series of air-stream surveys were made n the vicinity of the empennage of a 1/8.33-scale powered model of the XF-12 airplane in the Langley 19-foot pressure tunnel. Surveys of the vortical-tail region were made through a range of yaw angles of plus or minus 20 degrees at a high and low angle of attack. The horizontal-tail surveys were made over a fairly wide range of angles of attack at zero degrees yaw. Several power and flap conditions were investigated. The results are presented in the form a dynamic pressure ratios, sidewash angles, and downwash angles plotted against vertical distance from the fuselage center line. The results of the investigation indicate that a vertical tail located in a conventional position would be in a field of flow where the dynamic pressure ratios at the horizontal tail to be increased; for equal lift coefficients, the effect of power or flap deflection on the direction of flow at any particular point in the region of the horizontal tail is small.
Date: April 8, 1947
Creator: Foster, Gerald V.
Partner: UNT Libraries Government Documents Department

Altitude cooling investigation of the R-2800-21 engine in the P-47G airplane II : investigation of the engine & airplane variables affecting the cylinder temperature distribution

Description: The data obtained from cooling tests of an R-2800-21 engine installed in a p-47G airplane were studied to determine which engine and airplane operation variables were mainly responsible for the extremely uneven temperature distribution among the 18 engine cylinders obtained at the medium and high engine-power conditions. The tests consisted of flights at altitudes from 5000 to 35,000 feet for the normal range of engine and airplane operation. The results of the study showed that a flow condition in the induction system associated with the wide-open throttle position, which affected either the fuel air or charge distribution, was primarily responsible for the uneven temperature distribution. For the range of fuel-air ratios tested (0.080 to 0.102), the temperature distribution remained essentially unchanged. The individual effects of thrust-axis inclination, cowl-flap opening, and quantity of auxiliary air were found to be secondary in importance. At low angles of throttle opening, engine speed was found to have little effect on the temperature pattern.
Date: October 9, 1946
Creator: Pesman, Gerard J & Kaufman, Samuel J
Partner: UNT Libraries Government Documents Department

Altitude cooling investigation of the R-2800-21 engine in the P-47g airplane III : individual-cylinder temperature reduction by means of intake-pipe throttle and by coolant injection

Description: Flight tests were conducted on a R-2800-21 engine in the P-47G airplane to determine the effect on the wall temperatures of cylinder 10 of throttling the charge in the intake pipe and of injecting a water-ethanol coolant into the intake pipe. Cylinder 10 was chosen for this investigation because it runs abnormally hot (head temperatures of the order of 45 F higher than those of the next hottest cylinder) at the medium and high-power conditions. Tests with interchanged cylinders showed that the excessive temperatures of cylinder 10 were inherent in the cylinder location and were not due to the mechanical condition of the cylinder assembly. Throttling the charge in the intake pipe is a simpler method than coolant injection into the intake pipe particularly when only one cylinder is considerably hotter than any other. Coolant injection into the individual cylinders is a more efficient method than throttling in the intake pipe and is warranted when several cylinders are to be cooled or when parts of the complex equipment required are already available.
Date: October 9, 1946
Creator: Bell, E Barton; Valerino, Michael F & Manganiello, Eugene J
Partner: UNT Libraries Government Documents Department

Altitude Cooling Investigation of the R-2800-21 Engine in the P-47G Airplane. IV - Engine Cooling-Air Pressure Distribution

Description: A study of the data obtained in a flight investigation of an R-2800-21 engine in a P-47G airplane was made to determine the effect of the flight variables on the engine cooling-air pressure distribution. The investigation consisted of level flights at altitudes from 5000 to 35,000 feet for the normal range of engine and airplane operation. The data showed that the average engine front pressures ranged from 0.73 to 0.82 of the impact pressure (velocity head). The average engine rear pressures ranged from 0.50 to 0.55 of the impact pressure for closed cowl flaps and from 0.10 to 0.20 for full-open cowl flaps. In general, the highest front pressures were obtained at the bottom of the engine. The rear pressures for the rear-row cylinders were .lower and the pressure drops correspondingly higher than for the front-row cylinders. The rear-pressure distribution was materially affected by cowl-flap position in that the differences between the rear pressures of the front-row and rear-row cylinders markedly increased as the cowl flaps were opened. For full-open cowl flaps, the pressure drops across the rear-row cylinders were in the order of 0.2 of the impact pressure greater than across the front-row cylinders. Propeller speed and altitude had little effect on the -coolingair pressure distribution, Increase in angle of inclination of the thrust axis decreased the front ?pressures for the cylinders at the top of the engine and increased them for the cylinders at the bottom of the engine. As more auxiliary air was taken from the engine cowling, the front pressures and, to a lesser extent, the rear pressures for the cylinders at the bottom of the engine decreased. No correlation existed between the cooling-air pressure-drop distribution and the cylinder-temperature distribution.
Date: January 7, 1947
Creator: Kaufman, Samuel J.; Staudt, Robert C. & Valerino, Michael F.
Partner: UNT Libraries Government Documents Department

Altitude performance and operational characteristics of 29-inch-diameter tail-pipe burner with several fuel systems and flame holders on J35 turbojet engine

Description: An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning has been conducted in the NACA Lewis altitude wind tunnel. Several fuel systems and flame holders were investigated in a 29-inch-diameter tail-pipe burner to determine the effect of fuel distribution and flame-holder design on tail-pipe-burner performance and operational characteristics over a range of simulated flight conditions. At an altitude of 5000 feet, the type of flame holder used had only a slight effect on the combustion efficiency. As the altitude was increased, the decrease in peak combustion efficiency became more rapid as the blocking area of the flame holder was reduced. At all altitudes investigated, an improvement in the uniformity of the radial distribution of fuel and air slightly increased the peak combustion efficiencies and shifted the peak combustion efficiency to higher tail-pipe fuel-air ratios. The use of an internal cooling liner extending the full length of the tail-pipe combustion chamber provided adequate shell cooling at all flight conditions investigated.
Date: November 8, 1949
Creator: Conrad, E William & Prince, William R
Partner: UNT Libraries Government Documents Department

Altitude Performance of AN-F-58 Fuels in British Rolls-Royce Nene Single Combustor

Description: An investigation was conducted with a single combustor from a British Rolls-Royce Nene turbojet engine to determine the altitude performance characteristics of AN-F-58 fuels. Three fuel blends conforming to AN-F-58 specifications were prepared in order to determine the influence of fuel boiling temperatures and aromatic content on combustion efficiencies and altitude operational limits. The performance of the three AN-F-58 fuels was compared in the range of altitudes from sea level to 65,000 feet, engine speeds from 40- to 100- percent normal rated, and flight Mach numbers of 0.0 and 0.6. Similar information was obtained for AN-F-32 fuel at a flight Mach number of 0.0.
Date: July 8, 1949
Creator: Cook, William P. & Koch, Richard G.
Partner: UNT Libraries Government Documents Department

Altitude-Test-Chamber Investigation of a Solar Afterburner on the 24C Engine I - Operational Characteristics and Altitude Limits

Description: An altitude-test-chamber investigation was conducted to determine the operational characteristics and altitude blow-out limits of a Solar afterburner in a 24C engine. At rated engine speed and maximum permissible turbine-discharge temperature, the altitude limit as determined by combustion blow-out occurred as a band of unstable operation of about 8000 feet altitude in width with maximum altitude limits from 32,000 feet at a Mach number of 0.3 to about 42,000 feet at a Mach number of 1.0. The maximum fuel-air ratio of the afterburner, as limited by maximum permissible turbine-discharge gas temperatures at rated engine speed, varied between 0.0295 and 0.0380 over a range of flight Mach numbers from 0.25 to 1.0 and at altitudes of 20,000 and 30,000 feet. Over this range of operating conditions, the fuel-air ratio at which lean blow-out occurred was from 10 to 19 percent below these maximum fuel-air ratios. Combustion was very smooth and uniform during operation; however, ignition of the burner was very difficult throughout the investigation. A failure of the flame holder after 12 hours and 15 minutes of afterburner operation resulted in termination of the investigation.
Date: July 6, 1948
Partner: UNT Libraries Government Documents Department

Altitude-Test-Chamber Investigation of Mcdonnell Afterburner on J34 Engine

Description: An altitude-test-chamber investigation was conducted to determine the operational and performance characteristics of a McDonnell afterburner with a fixed-area exhaust nozzle on a J34 engine. At rated engine speed, the altitude limit, as determined by combustion blow-out, occurred as a band of unstable operation of about 6000-foot altitude in width with minimum altitude limits from 31,000 feet at a simulated flight Mach number of 0.40 to about 45,500 feet at a simulated flight Mach number of 1.00. Considerable difficulty was experienced in attempting to establish or maintain balanced-cycle engine operation at altitudes above 36,000 feet. The fuel-air ratio for balanced-cycle operation and lean blowout of the afterburner, the augmented-thrust ratio, the total specific fuel consumption, and the afterburner combustion efficiency for balanced-cycle operation are summarized in a table. Satisfactory afterburner ignition was obtained over a range of flight Mach Numbers from 0.32 to 0.60 at altitudes from 10,000 to 30,000 and engine speeds from 10,000 to 12,500 rpm.
Date: April 18, 1949
Creator: Reller, John O. & Dowman, Harry W.
Partner: UNT Libraries Government Documents Department

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine, 3, Analysis of Combustion-Chamber Performance

Description: Combustion chamber performance properties of a 3000-pound-thrust axial-flow turbojet engine were determined. Data are presented for a range of simulated altitudes from 15,000 to 45,0000 feet and a range of Mach numbers from 0.23 to 1.05 for various modifications of the engine.
Date: August 23, 1948
Creator: Campbell, Carl E.
Partner: UNT Libraries Government Documents Department

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine, 4, Operational Characteristics

Description: An investigation was conducted to evaluate the operational characteristics of a 3000 pound thrust axial flow turbojet engine over a range of simulated altitudes from 2000 to 50,000 feet and simulated flight Mach numbers from 0 to 1.04 throughout the operable range of engine speeds. Engine operating range, acceleration, deceleration, starting, altitude, and flight Mach number compensation of the fuel control system, and operation of the lubrication system at high and low ambient air temperatures were evaluated.
Date: August 23, 1948
Creator: Hawkins, W. Kent & Meyer, Carl L.
Partner: UNT Libraries Government Documents Department

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine, 6, Analysis of Effects of Inlet Pressure Losses

Description: The losses in the inlet air ducts, the diffusers, and the de-icing equipment associated with turbojet engine installations cause a reduction in the total pressure at the inlet of the engine and result in reduced thrust and increased specific fuel consumption. An analytical evaluation of the effects of inlet losses on the net thrust and the fuel economy of a 3000-pound-thrust axial flow turbojet engine with a two-stage turbine is presented. The analysis is based on engine performance characteristics that were determined from experiments in the NACA Cleveland altitude wind tunnel. The experimental investigation did not include tests in which inlet losses were systematically varied, but the effects of these losses can be accurately estimated from the experimentally determined performance characteristics of the engine.
Date: 1948
Creator: Sanders, Newell D. & Palasics, John
Partner: UNT Libraries Government Documents Department

Altitude-Wind-Tunnel Investigation of a 3000-Pound-Thrust Axial-Flow Turbojet Engine, 7, Pressure and Temperature Distributions

Description: Temperature and pressure distributions for an original and modified 3000 pound thrust axial flow turbojet engine were investigated. Data are included for a range of simulated altitudes from 5000 to 45000 feet, Mach numbers from 0.24 to 1.08, and corrected engine speeds from 10,550 to 13,359 rpm.
Date: December 10, 1948
Creator: Saari, Martin J. & Prince, William R.
Partner: UNT Libraries Government Documents Department