National Advisory Committee for Aeronautics (NACA) - 37 Matching Results

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Longitudinal-stability investigation of high-lift and stall-control devices on a 52 degree sweptback wing with and without fuselage and horizontal tail at a Reynolds number of 6.8 x 10(exp 6).

Description: Contains low-speed longitudinal stability characteristics of a 52 degree sweptback wing of aspect ratio 2.88, taper ratio 0.625, and NACA 64 (sub 1)-112 airfoil sections normal to the 0.282-chord line, in combination with split flaps, leading-edge flaps, and upper-surface fences. Low-wing and midwing-fuselage aerodynamic characteristics are presented with and without a horizontal tail at various vertical locations. Tests were conducted at a Reynolds number of 6.8 x 10(exp 6).
Date: December 20, 1948
Creator: Foster, Gerald V. & Fitzpatrick, James E.
Partner: UNT Libraries Government Documents Department

Determination of coupled modes and frequencies of swept wings by use of power series

Description: From Summary: "A solution is presented for the coupled modes and frequencies of swept wings mounted on a fuselage. The energy method is used in conjunction with power series to obtain the characteristic equations for both symmetrical and asymmetrical vibration. A numerical example which is susceptible to exact solution is presented, and the results for the exact solution and the solution presented in this paper show excellent agreement."
Date: October 20, 1947
Creator: Anderson, Roger A.
Partner: UNT Libraries Government Documents Department

Drag measurements of a 34 degree swept-forward and swept-back NACA 65-009 airfoil of aspect ratio 2.7 as determined by flight tests at supersonic speeds

Description: Report presenting the results of flight testing to determine the zero-lift drag of an NACA 65-009 airfoil at a specified aspect ratio. The results are compared to previous testing of unswept and swept-back arrangements. The swept-forward and swept-back airfoils were found to produce lower values of zero-drag lift than the unswept airfoil.
Date: February 20, 1947
Creator: Alexander, Sidney R.
Partner: UNT Libraries Government Documents Department

Measurements of aileron effectiveness of Bell X-1 airplane up to a Mach number of 0.82

Description: From Summary: "Abrupt, rudder-fixed aileron rolls have been made with the Bell X-1 airplane having a 10-percent-thick wing in glides to a Mach number of 0.82 at about 30,000 feet pressure altitude. Aileron movements were between one-fourth and one-half of full deflection. These aileron rolls indicate that Mach number has little effect on the aileron effectiveness up to a Mach number of 0.82."
Date: June 20, 1949
Creator: Drake, Hubert M.
Partner: UNT Libraries Government Documents Department

A free-flight technique for measuring damping in roll by use of rocket-powered models and some initial results for rectangular wings

Description: Report presenting a simplified method for obtaining free-flight measurements of damping in roll through the use of rocket-powered models. Initial configurations have been tested for a range of Mach numbers. Results regarding the rolling velocity with two different airfoil sections and damping-in-roll coefficient are provided.
Date: December 20, 1949
Creator: Edmondson, James L. & Sanders, E. Claude, Jr.
Partner: UNT Libraries Government Documents Department

Measurements of the effects of thickness ratio and aspect ratio on the drag of rectangular-plan-form airfoils at transonic speeds

Description: Report presenting testing conducted on two airfoils from a series of rectangular-plan-form airfoils of aspect ratios 7.6 and 5.1 and with NACA 65-006, 65-009, and 65-012 sections using the free-fall method. Results regarding the time histories, ground-velocity data, airfoil drag measurements, and drag coefficients are provided.
Date: June 20, 1947
Creator: Thompson, Jim Rogers & Mathews, Charles W.
Partner: UNT Libraries Government Documents Department

Effect of a pilot's canopy on the drag of an NACA RM-2 drag research model in flight at transonic and supersonic speeds

Description: Report presenting data from two experiments. One used the NACA RM-2 drag research model equipped with a pilot's canopy to determine the effect on aerodynamics. The other was conducted with the same configuration and returned similar results.
Date: April 20, 1948
Creator: Purser, Paul E.
Partner: UNT Libraries Government Documents Department

Preliminary Tests in the Supersonic Sphere

Description: This report presents preliminary data obtained in the Langley supersonic sphere. The supersonic sphere is essentially a whirling mechanism enclosed in a steel shell which can be filled with either air or Freon gas. The test models for two-dimensional study are of propeller form having the same plan form and diameter but varying only in the airfoil shape and thickness ratio. Torque coefficients for the 16-006, 65-110, and the 15 percent thick ellipse models are presented, as well as pressure distributions on a circular-arc supersonic airfoil section having a maximum thickness of 10 percent chord at the 1/3-chord position. Torque coefficients were measured in both Freon and air on the 15 percent thick ellipse, and the data obtained in air and Freon are found to be in close agreement. The torque coefficients for the three previously mentioned models showed large differences in magnitude at tip Mach numbers above 1, the model with the thickest airfoil section having the largest torque coefficient. Pressure distribution on the previously mentioned circular-arc airfoil section are presented at Mach numbers of 0.69, 1.26, and 1.42. At Mach numbers of 1.26 and 1.42 the test section is in the mixed flow region where both subsonic and supersonic speeds occur on the airfoil. No adequate theory has been developed for this condition of mixed flow, but the experimental data have been compared with values of pressure based on Ackeret's theory. The experimental data obtained at a Mach number of 1.26 on the rear portion of the airfoil section agree fairly well with the values calculated by Ackeret's theory. At a Mach number of 1.42 a larger percentage of the airfoil is in supersonic flow, and the experimental data for the entire airfoil agree fairly well with the values obtained using Ackeret's theory.
Date: January 20, 1947
Creator: Baker, John E.
Partner: UNT Libraries Government Documents Department

Summary of Section Data on Trailing-Edge High-Lift Devices

Description: A summary has been made of available data on the characteristics of airfoil sections with trailing-edge high-lift devices. Data for plain, split, and slotted flaps are collected and analyzed. The effects of each of the variables involved in the design of the various types of flap are examined and, in cases where sufficient data are given, optimum configurations are deduced. Wherever possible, the effects of airfoil section, Reynolds number, and leading-edge roughness are shown. For single and double slotted flaps, where a great mass of unrelated date are available, maximum lift coefficients of a large number of configurations are presented in tables.
Date: August 20, 1948
Partner: UNT Libraries Government Documents Department

Flight Investigation of the Effects of Ice on an I-16 Jet-Propulsion Engine

Description: A flight investigation of an I-16 jet propulsion engine installed in the waist compartment of a B-24M airplane was made to determine the effect of induction-system icing on the performance of the engine. Flights were made at inlet-air temperatures of 15 deg, 20 deg., and 25 F, an indicated airspeed of 180 miles per hour, jet-engine speeds of 13,000 and 15,000 rpm, liquid-water contents of approximately 0.3 to 0.5 gram per cubic meter, and an average water droplet size of approximately 50 microns. Under the most severe icing conditions obtained, ice formed on the screen over the front inlet to the compressor and obstructed about 70 percent of the front-inlet area. The thrust was thereby reduced 13.5 percent, the specific fuel consumption increased 17 percent, and the tail-pipe temperature increased 82 F. No icing of the rear compressor-inlet screen was encountered.
Date: January 20, 1947
Creator: Pragliola, Philip C. & Werner, Milton
Partner: UNT Libraries Government Documents Department

Effect of Exhaust Pressure on the Cooling Characteristics of a Liquid-Cooled Engine

Description: Data for a liquid-cooled engine with a displacement volume of 1710 cubic inches were analyzed to determine the effect of exhaust pressure on the engine cooling characteristics. The data covered a range of exhaust pressures from 7 to 62 inches of mercury absolute, inlet-manifold pressures from 30 to 50 inches of mercury absolute, engine speeds from 1600 to 3000 rpm, and fuel-air ratios from 0.063 to 0.100. The effect of exhaust pressure on engine cooling was satisfactorily incorporated in the NACA cooling-correlation method as a variation in effective gas temperature with exhaust pressure. Large variations of cylinder-head temperature with exhaust pressure were obtained for operation at constant charge flow. At a constant charge flow of 2 pounds per second (approximately 1000 bhp) and a fuel-air ratio of 0.085, an increase in exhaust pressure from 10 to 60 inches of mercury absolute resulted in an increase of 40 F in average cylinder-head temperature. For operation at constant engine speed and inlet-manifold pressure and variable exhaust pressure (variable charge flow), however, the effect of exhaust pressure on cylinder-head temperature is small. For example, at an inlet-manifold pressure of 40 inches of mercury absolute, an engine speed of 2400 rpm.- and a fuel-air ratio of 0.085, the average cylinder-head temperature was about the same at exhaust pressures of 10 and 60 inches of,mercury absolute; a rise and a subsequent decrease of about 70 occurred between these extremes.
Date: January 20, 1947
Creator: Doyle, Ronald B. & Desmon, Leland G.
Partner: UNT Libraries Government Documents Department

Performance of J33-A-23 Turbojet-Engine Compressor, Part 1, Over-All Performance Characteristics of Compressor with 17-Blade Impeller

Description: The production-model 333-A-23 turbojet-engine compressor with a 17-blade impeller was operated at ambient and 0 F inlet temperatures and at inlet pressures of 14 and 5 inches mercury absolute for equivalent impeller speeds from 6000 to 12,750 rpm. The results of this investigation are compared with those of the 533-A-21 compressor. At the design equivalent speed of 11,750 rpm the maximum pressure ratio was 4.39. This occurred at the surge point at which the equivalent weight flow was 80.8 pounds per second, ana the adiabatic temperature-rise efficiency was 0.757. The maximum flow at the design equivalent speed was 88.0 pounds per second. The maximum adiabatic temperature-rise efficiency of 0.799 was obtained at an equivalent speed of 10,000 rpm, and equivalent weight flow of 62.9 pounds per second, and a pressure ratio of 3.20. At the maximum equivalent speed investigated (12,750 rpm), a peak pressure ratio of 4.90 was attained at an equivalent weight flow of 85.4 pounds per second and an efficiency of 0.680.
Date: July 20, 1948
Creator: Beede, William L. & Kottas, Harry
Partner: UNT Libraries Government Documents Department

Performance of the Modified V-1710-93 Engine-Stage Supercharger with a Constant-Area Vaneless Diffuser

Description: As part of an investigation to increase the power output of the V-1710-93 engine at altitude, the engine-stage supercharger was combined with a constant-area vaneless diffuser designed to improve the performance of the engine-stage supercharger at the rated engine operating point. The performance of the modified supercharger was investigated in a variable-component supercharger test rig and compared with that of the standard supercharger with an 8-vaned diffuser. A separate evaluation of the component efficiencies and a study of the flow characteristics of the modified supercharger was made possible by internal diffuser instrumentation. At the volume flow required by the engine for rated operating conditions, the modified supercharger increased the over-all adiabatic efficiency 0.05 and the over-all pressure coefficient 0.035. Furthermore, the capacity of the engine-stage supercharger was increased by replacing the standard 8-vaned diffuser with the vaneless diffuser. The peak over-all adiabatic efficiency for the modified supercharger, however, was 0.05 to 0.07 lower than that of the standard unit over the range of tip speeds investigated. The improved performance of the modified supercharger at rated engine operating conditions resulted from a shift of the point of peak adiabatic efficiency and pressure coefficient of the standard supercharger to a higher flow. The energy loss through the vaneless diffuser was found to be small. Because of the restricted diffuser diameter, however, diffusion was inadequate, which resulted in a relatively small static-pressure rise through the diffuser, high diffuser-exit velocities, and excessive collector-case losses.
Date: December 20, 1946
Creator: Douglas, John E. & Schwartz, Irving R.
Partner: UNT Libraries Government Documents Department

Effects of Induction-System Icing on Aircraft-Engine Operating Characteristics

Description: An investigation was conducted on a multicylinder aircraft engine on a dynamometer stand to determine the effect of induction-system icing on engine operating characteristics and to compare the results with those of a previous laboratory investigation in which only the carburetor and the engine-stage supercharger assembly from the engine were used. The experiments were conducted at simulated glide power, low cruise power, and normal rated power through a range of humidity ratios and air temperatures at approximately sea-level pressure. Induction-system icing was found to occur within approximately the same limits as those established by the previous laboratory investigation after making suitable allowances for the difference in fuel volatility and throttle angles. Rough operation of the engine was experienced when ice caused a marked reduction in the air flow. Photographs of typical ice formations from this investigation indicate close similarity to icing previously observed in the laboratory.
Date: January 20, 1947
Creator: Stevens, Howard C., Jr.
Partner: UNT Libraries Government Documents Department