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NACA: University Conference on Aerodynamics, A Compilation of the Papers Presented
This document contains reproductions of the technical papers presented at the NACA - University Conference on Aerodynamics held at the Langley Aeronautical Laboratory on June 21, 22, and 23, 1948. The conference was held in recognition of the difficulties, imposed by security restrictions, in keeping abreast of the rapid advances in aerodynamics. The papers were prepared to review the status of a number of fields of interest, to summarize the more important wartime advances that are no longer classified, and to orient reference material for further study. The papers in this document are in the same form in which they were presented at the conference so that distribution of them might be prompt. The original presentation and this record are considered as complementary to, rather than as substitutes for, the Committee?s system of complete and formal reports.
A profile-drag investigation in flight on an experimental fighter-type airplane the North American XP-51
No Description Available.
Wind-tunnel investigation of several factors affecting the performance of a high-speed pursuit airplane with air-cooled radial engine
No Description Available.
Experimental investigation of a new type of low-drag wing-nacelle combination
No Description Available.
Model tests of a wing-duct system for auxiliary air supply
No Description Available.
Turbine Engines for High-Speed Flight
This analysis investigates the application of gas turbine engines at a cruise Mach number of 4.
AVA monographs. B: Boundary layer
No Description Available.
NACA conference on Aerodynamic Problems of Transonic Airplane Design
No Description Available.
Analysis of the High-Altitude Cooling of the Ranger SGV-770 D-4 Engine in the Bell XP-77 Airplane
No abstract available.
Annual Report of the National Advisory Committee for Aeronautics (42nd), Administrative Report without Technical Reports
No abstract available.
High-altitude cooling. III : Radiators
No Description Available.
Tests of a heated low-drag airfoil
No Description Available.
Investigation of Wingless Missile Configurations with Folding Controls and Low-Aspect-Ratio Stabilizing Surfaces
Wind tunnel tests of wingless low aspect ratio folding control missile configurations.
A method for calculating the lift and center of pressure of wing-body-tail combinations at subsonic, transonic, and supersonic speeds
No Description Available.
Method for estimating pitching-moment interference of wing-body combinations at supersonic speed
No Description Available.
Preliminary Results from Free-Jet Tests of a 48-Inch-Diameter Ram-Jet Combustor with an Annular-Piloted Baffle-Type Flameholder
A ram-jet engine with an experimental 48-inch-diameter combustor was investigated in a free-jet facility. The combustor design comprised a large-volume annular pilot region and an array of sloping baffle- or gutter-type flameholders. The combustor was intended to operate at a fuel-air ratio of about 0.037. To promote combustion efficiency at such low fuel-air ratios, a divided-flow system was employed which bypassed a portion of the engine air around the combustion region. Three combustor lengths, three lengths of the shroud which separated the bypass air from the burning stream, and four fuel-distribution systems were investigated over a range of fuel-air ratios from 0.025 to 0.055 and a range of engine air flows from 40 to 110 pounds per second (combustor-outlet total pressures from 500 t o 1800 lb/sq ft abs). The highest efficiencies were obtained with a combustor length of 78 inches and a shroud length of 6 inches. At the lowest air flow, with combustor pressures of about 700 pounds per square foot absolute, a maximum efficiency of about 93 percent was obtained. The efficiency increased with combustor length, a typical increase being from 88 to 95 percent as the length increased from 60 to 96 inches. The length of the shroud separating the bypass air from the burning stream affected not only the efficiency level, but also the fuel-air ratio at which the maximum efficiency occurred. In general, a longer shroud caused the maximum efficiency to occur at lower f'uel-air ratios. Highest efficiencies usually resulted from the use of a fuel-injection system giving a uniform fuel profile. The efficiency at low fuel-air ratios could be considerably improved by the use of a radially nonuniform fuel profile which concentrated the fuel towards the outermost portion of the burning stream The total-pressure ratio across the combustor was about 0.86 at the ...
Preliminary Data on Rain Deflection from Aircraft Windshields by Means of High-Velocity Jet-Air Blast
A preliminary experimental investigation is being conducted to determine the feasibility of preventing rain from impinging on aircraft windshields by means of high-velocity jet-air blast. The results indicate that rain deflection by jet blast appears feasible for flight speeds comparable with landing and take-off speeds of interceptor-type jet aircraft; however, attainment of good visibility through the mist generated by raindrop breakup presents a problem. For the simulated windshield and the lower windshield angles used in the investigation, air-flow rates of the order of 3.3 pounds per minute of unheated air per inch of windshield span were required for adequate rain deflection at a free-stream velocity of 135 miles per hour. A method has been devised whereby it is possible to produce large-diameter water drops (1000 to 1500 p.) in a moving air stream, without breakup, at speeds in excess of 175 miles per hour.
Flight Instrument for Measurement of Liquid-Water Content in Clouds at Temperatures Above and Below Freezing
A principle formerly used in an instrument for cloud detection was further investigated to provide a simple and rapid means for measuring the liquid-water content of clouds at temperatures above and below freezing. The instrument consists of a small cylindrical element so operated at high surface temperatures that the impingement of cloud droplets creates a significant drop in the surface temperature. ? The instrument is sensitive to a wide range of liquid-water content and was calibrated at one set of fixed conditions against rotating multicylinder measurements. The limited conditions of the calibration Included an air temperature of 20 F, an air velocity of 175 miles per hour, and a surface temperature in clear air of 475 F. The results obtained from experiments conducted with the instrument indicate that the principle can be used for measurements in clouds at temperatures above and below freezing. Calibrations for ranges of airspeed, air temperature, and air density will be necessary to adapt the Instrument for general flight use.
Impingement of Water Droplets on an NACA 65(sub 1) -212 Airfoil at an Angle of Attack of 4 Deg
The trajectories of droplets in the air flowing past an NACA 651-212 airfoil at an angle of attack of 40 were determined. The collection efficiency, the area of droplet impingement, and the rate of droplet impingement were calculated from the trajectories and are presented herein.
Investigation of Porous Gas-Heated Leading-Edge Section for Icing Protection of a Delta Wing
A tip section of a delta wing having an NACA 0004-65 airfoil section and a 600 leading-edge sweepback was equipped with a porous leading-edge section through which hot gas was 'bled for anti-icing. Heating rates for anti-icing were determined for a wide range of icing conditions. The effects of gas flow through the porous leading-edge section on airfoil pressure distribution and drag in dry air were investigated. The drag increase caused by an ice formation on the unheated airfoil was measured for several icing conditions. Experimental porous surface- to free-stream convective heat-transfer coefficients were obtained in dry air and compared with theory. Adequate icing protection was obtained at all icing conditions investigated. Savings in total gas-flow rate up to 42 percent may be obtained with no loss in anti-icing effectiveness by sealing half the upper-surface porous area. Gas flow through the leading-edge section had no appreciable effect on airfoil pressure distribution. The airfoil section drag increased slightly (5-percent average) with gas flow through the porous surface. A heavy glaze-ice formation produced after 10 minutes of icing caused an increase in section drag coefficient of 240 percent. Experimental convective heat-transfer coefficients obtained with hot-gas flow through the porous area in dry air and turbulent flow were 20 to 30 percent lower than the theoretical values for a solid surface under similar conditions. The transition region from laminar to turbulent flow moved forward as the ratio of gas velocity through the porous surface to air-stream velocity was increased.
Investigation of Power Requirements for Ice Prevention and Cyclical De-Icing of Inlet Guide Vanes with Internal Electric Heaters
An investigation was conducted to determine the electric power requirements necessary for ice protection of inlet guide vanes by continuous heating and by cyclical de-icing. Data are presented to show the effect of ambient-air temperature, liquid-water content, air velocity, heat-on period, and cycle times on the power requirements for these two methods of ice protection. The results showed that for a hypothetical engine using 28 inlet guide vanes under similar icing conditions, cyclical de-icing can provide a total power saving as high as 79 percent over that required for continuous heating. Heat-on periods in the order of 10 seconds with a cycle ratio of about 1:7 resulted in the best over-all performance with respect to total power requirements and aerodynamic losses during the heat-off period. Power requirements reported herein may be reduced by as much as 25 percent by achieving a more uniform surface-temperature distribution. A parameter in terms of engine mass flow, vane size, vane surface temperature, and the icing conditions ahead of the inlet guide vanes.was developed by which an extension of the experimental data to icing conditions and inlet guide vanes, other than those investigated was possible.
Tracking Performance of a Swept-wing Fighter With a Directortype Radar Fire-control System and Scope Presentation
Tracking performance of f-86d aircraft with radar fire-control system.
Transonic wind tunnel tests of the launch, jettison, and longitudinal characteristics of an airplane and missile model combination
No Description Available.
Wind-tunnel Investigation at Subsonic and Supersonic Speeds of a Model of a Tailless Fighter Airplane Employing a Low-aspect-ratio Swept-back Wing-stability and Control
Subsonic and supersonic wind tunnel study of low aspect ratio sweptback wing tailless fighter aircraft model/stability and control.
Wind-tunnel investigation of a 1/6-scale model of the Bumblebee XPM missile at high subsonic speeds
No Description Available.
Wind-Tunnel Investigation of a 1/60-Scale Model of the Republic MX-1554 Airplane at a Mach Number of 2.85
No Description Available.
A wind tunnel investigation of several wingless missile configurations at supersonic speeds
No Description Available.
Wind-Tunnel Investigation of the Aerodynamic Characteristics of a 0.07-Scale Model of the North American MX-770 Missile
No Description Available.
Wind-Tunnel Investigation of the Drag and Lateral-Stability Characteristics of a 1/22-Scale Model of a Bomber Airplane Employing a Low-Aspect-Ratio Triangular Wing
No Description Available.
Wind-tunnel measurements at subsonic speeds of the static and dynamic-rotary stability derivatives of a triangular-wing airplane model having a triangular vertical tail
No Description Available.
Wind-tunnel Tests of the Static Longitudinal Characteristics at Low Speed of a Swept-wing Airplane With Blowing Flaps and Leading-edge Slats
Low speed aerodynamic characteristics of swept- wing aircraft with blowing flaps and leading-edge slats.
Wing-body combinations with wings of very low aspect ratio at supersonic speeds
Wing-body combinations with wings of very low aspect ratio at supersonic speeds.
Boundary-layer-transition measurements in full-scale flight
No Description Available.
Boundary-Layer-Transition Measurements in Full-Scale Flight
Chemical sublimation has been employed for boundary-layer-flow visualization on the wings of a supersonic fighter airplane in level flight at speeds near a Mach number of 2.0. The tests have shown that laminar flow can be obtained over extensive areas of the wing with practical wing-surface conditions. In addition to the flow visualization tests, a method of continuously monitoring the conditions of the boundary layer has been applied to flight testing, using heated temperature resistance gages installed in a Fiberglas "glove" installation on one wing. Tests were conducted at speeds from a Mach number of 1.2 to a Mach number of 2.0, at altitudes from 35,000 feet to 56,000 feet. Data obtained at all angles of attack, from near 0 deg to near 10 deg, have shown that the maximum transition Reynolds number on the upper surface of the wing varies from about 2.5 x 10(exp 6) at a Mach number of 1.2 to about 4 x 10(exp 6) at a Mach number of 2.0. On the lower surface, the maximum transition Reynolds number varies from about 2 x 10(exp 6) at a Mach number of 1.2 to about 8 x 10(exp 6) at a Mach number of 2.0.
Altitude performance of a 20-inch-diameter ram-jet engine investigated in a free-jet facility at Mach number 3.0
No Description Available.
Comparison and evaluation of two model techniques used in predicting bomb-release motions
No Description Available.
Pressure Distributions and Aerodynamic Characteristics of Several Spoiler-type Controls on a Trapezoidal Wing at Mach Numbers of 1.61 and 2.01
Pressure distributions and aerodynamic characteristics of spoiler-type controls on trapezoidal wing at supersonic speed.
Aerodynamic characteristics of a 0.04956-scale model of the Convair F-102A airplane at Mach numbers of 1.41, 1.61, and 2.01
Tests have been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.41, 1.61, and 2.01 of various arrangements of a 0.04956-scale model of the Convair F-102A airplane with faired inlets. Tests made of the model equipped with a plain wing, a wing with 6.4 percent conical camber, and a wing with 15 percent conical camber. Body modifications including an extended nose, a modified canopy, and extended afterbody fillets were evaluated. In addition, the effects of a revised vertical tail and two different ventral fins were determined. The results indicated that the use of cambered wings resulted in lower drag in the lift-coefficient range above 0.2. This range, however, is above that which would generally be required for level flight; hence, the usefulness of camber might be confined to increased maneuverability at the higher lifts while its use may be detrimental to the high-speed (low-lift) capabilities.
Transonic flight evaluation of the effects of fuselage extension and indentation on the drag of a 60 deg delta wing interceptor airplane
No Description Available.
Analysis of flight-determined and predicted effects of flexibility on the steady-state wing loads of the B-52 airplane
Flight-determined and predicted effects of flexibility on steady-state wing loads of B-52 airplane.
Induction system characteristics and engine surge occurrence for two fighter-type airplanes
No Description Available.
In-flight Gains Realized by Modifying a Twin Side-inlet Induction System
Modification of twin side-inlet induction system and in-flight gains.
Power-off tests of the Northrop N9M-2 tailless airplane in the 40- by 80-foot wind tunnel
No Description Available.
Heat Transfer Measured in Free Flight on a Slightly Blunted 25 deg Cone-Cylinder-Flare Configuration at Mach Numbers up to 9.89
Skin temperature and surface pressure of blunted cone-cylinder-flare configuration free flight test vehicle to hypersonic speeds.
Model ditching investigations of three airplanes equipped with hydro-skis
No Description Available.
A note on the ability to predict transonic drag-rise changes due to model modifications
Wind tunnel test to establish validity of Fourier analysis results for transonic drag rise changes due to model modifications.
The Origin and Distribution of Supersonic Store Interference From Measurement of Individual Forces on Several Wing-fuselage-store Configurations. Ii - Swept-wing Heavy-bomber Configuration With Large Store Nacelle . Lateral Forces and Pitching Moments, Mach Number, 1.61
Effect of large store /nacelle/ on a swept-wing heavy bomber - lateral force and pitching moments.
The Origin and Distribution of Supersonic Store Interference From Measurement of Individual Forces on Several Wing-fuselagestore Configurations. 1.-swept-wing Heavy-bomber Configuration With Large Store Nacelle . Lift and Drag, Mach Number, 1.61
Supersonic store interference - 1, swept-wing heavy bomber with large stores - lift & drag at mach 1.61.
The Origin and Distribution of Supersonic Store Interference From Measurement of Individual Forces on Several Wing-fuselagestore Configurations. Iii - Swept-wing Fighter-bomber Configuration With Large and Small Stores. Mach Number, 1.61
Origin of supersonic store interference from measurements of individual forces - swept-wing fighter-bomber with large & small stores.
Aerodynamics of Oscillating Control Surfaces at Transonic Speeds
Aerodynamic forces and moments of oscillating control surfaces at transonic speeds.