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A Preliminary Report on the MTR Crystal Spectrometer
Report regarding the installation and testing of a neutron crystal spectrometer. This was part of a slide show presentation.
Preliminary report on the problem of the atmosphere in relation to aeronautics
A report to the Weather Bureau, Washington DC, from the chairman of the Subcommittee on the Atmosphere in Relation to Aeronautics describing the activities accomplished and the proposal of work to be undertaken by the subcommittee.
Preliminary Report on the Results of Geobotanical Prospecting on the South Flank of Haystack Butte, McKinley County, New Mexico
Abstract: The absorber plant method of geobotanical prospecting was tested systematically over the bench formed by the Jurassic Todilto limestone on the south flank of Haystack Butte, McKinley County, N. Mex. This portion of the bench includes the largest known uranium ore body in limestone and most has been extensively drilled by private enterprise. Geobotanical prospecting was accomplished to provide control data. Comparison of the geobotanical anomalies with the available drill hole information from the mining companies and Atomic Energy Commission geologists have shown that the known ore occurrences would have been outlined by the results of the tree sampling. In addition some geobotanical anomalies are indicated in drilled areas in which ore was not reported and in areas not physically explored at the time of sampling. These anomalies may represent mineralized ground below ore grade or new ore deposits.
Preliminary Report on the Rimini Area, Jefferson City Quadrangle, Jefferson County and Lewis and Clark County, Montana
Abstract: A number of radioactivity anomalies and secondary uranium minerals have been found in the Rimini area near the northwestern margin of the Boulder batholith. Most of the anomalies are associated with chalcedonic vein zones that consist of one or more veins of cryptocrystalline and fine-grained quartz in silicified quartz monzonite and alaskite. Seventeen of the anomalies were found along veins that contain base and precious metals; nine were along veins in the vicinity of the village of Rimini from which there has been production of lead, silver, zinc, and gold.
Preliminary Report on the Uranium Occurrence of the Green Velvet Claims, Inyo County, California
Abstract: Several uranium occurrences have been discovered in the Coso lake beds, Owens Valley, California, along the west flank of the Coso Range.
Preliminary Report on the Uranium Occurrence of the Silver Lady Claim, Jaw Bone Mining District, Cross Mountain Quadrangle, Kern County, California
Abstract: An area of high anomalous radioactivity and visible uranium minerals is associated with a nearly vertical north 70° west trending shear zone.
Preliminary Report on Uranium-Bearing Deposits in Mohave County, Arizona
Abstract: Preliminary studies of the Wallapai Mining District and selected properties in the Maynard and Greenwood Mining Districts, Mohave County, Arizona, from January 8 to March 8, 1953, were made to determine the extent of uranium mineralization. All of the uranium properties examined are of the vein type and are believed to be of mesothermal origin. Brecciation and porosity of the veins appear to be controlling factors in the concentration of uranium minerals from the ore-bearing solutions. Although the uranium minerals present in these districts have not been specifically identified, they appear to be mostly primary with very minor occurrences of secondary products. One exception is the State mine in the Greenwood District, where secondary minerals predominate.
Preliminary Report on Uranium-Bearing Deposits of the Northern Boulder Batholith Region, Jefferson County, Montana
Introduction: This preliminary report is an attempt to correlate all of the available information having a bearing on the geology of the uraniferous deposits and to set forth such theories and recommendations as may be deduced at this stage.
Preliminary Report on Uranium-Bearing Pliocene (?) Rocks in the Split Rock Area, Central Wyoming
From abstract: A sequence of slightly radioactive Pliocene(?) rocks in the Split Rock area of central Wyoming is at least 600 feet thick, underlies an area of more than 25 square miles, and consists of alternating beds of tuffaceous shale, sandstone, and pumicite. Only a small part of the section has been examined. In addition to the uranium content of the strata, there has been secondary concentration of uranium in brecciated limestone similar to spring deposits that could be of Pliocene or Pleistocene age.
Preliminary Report on Uranium Deposits in the Gulf Coastal Plain, Southern Texas
Abstract: Concentration of secondary uranium minerals, some of which are commercially significant, have been found in three formations of Tertiary age in the Gulf Coastal Plain area of southern Texas: the Fayette sandstone of the Jackson formation, the Catahoula tuff, and the Oakville sandstone.
Preliminary Report on Uranium Deposits in the Wind River Basin, Wyoming
Abstract: Uranium was discovered in the Gas Hills of the Wind River Basin by Mr. Nail E. McNeice of Riverton, Wyoming while he was prospecting with a Geiger counter in September 1953. Field parties of the Atomic Energy Commission started work in the area in October 1953.
Preliminary Report on Uranium Occurrence, Silver King Claims, Tooele County, Utah
Abstract: Uranium was discovered on the Silver King claims in the fall of 1953. The claims are on the west flank of the Sheeprock Mountains in the eastern part of the Erickson mining district, Tooele County, Utah. Uraninite occurs in north- to northwest-trending copper-nickel-silver bearing fissure veins near the margin of a granitic stock of probable late Tertiary age. Sedimentary rocks in contact with the granite are chiefly dolomite and quartzite of Middle and Upper Ordovician age. Diamond drilling on this property did not disclose significant amounts of uranium; however, encouraging showings have been found by underground exploration by the owner.
A Preliminary Report on Uranium, Radium, and Vanadium
From Introduction: "This bulletin presents a summary of available information regarding the sources of uranium, radium, and vanadium, the methods used in treating the ores, and the uses of the the finished products. In particular the paper describes the ores found in the United States, giving especial attention to those characteristics of the ores and the conditions of their occurrence that affect mining and treatment."
Preliminary results from a limited investigation of the use of controls during service operational training with fighter airplanes
No Description Available.
Preliminary results from fatigue tests with reference to operational statistics
Simple elements were subjected to repeated loads of variable ampliture, chosen in such a way that they may be regarded as approximations to the operational loads (gust and maneuver) experienced by an airplane. The effect of varying some parameters was investigated briefly. Some discussion is given of the question whether a design according to current (1938 German) requirements for static strength is adequate from the fatigue point of view, and existing requirements on fatigue strength are compared,.
Preliminary results from flight measurements in gradual-turn maneuvers of the wing loads and the distribution of load among the components of a Boeing B-47A airplane
No Description Available.
Preliminary results from flow-field measurements around single and tandem rotors in the langley full-scale tunnel
No Description Available.
Preliminary results from free-jet tests of a 48-inch-diameter ram-jet combustor with an annular can-type flame holder
No Description Available.
Preliminary results from free-jet tests of a 48-inch-diameter ram-jet combustor with an annular can-type flame holder
Free jet tests of 48 inch diameter ramjet combustor with annular can-type flame holder.
Preliminary results from free jet tests of a 48-inch-diameter ram-jet combustor with an annular-piloted baffle-type flameholder
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 Results Obtained from Flight Test of a 1/7-Scale Rocket-Powered Model of the Grumman XF10F Airplane Configuration in the Swept-Wing Condition, TED No. NACA DE 354
A flight investigation of a 1/7-scale rocket-powered model of the XF10F Grumman XFl0F airplane in the swept-wing configuration has been made. The purpose of this test was to determine the static longitudinal stability, damping in pitch, and longitudinal control effectiveness of the airplane with the center of gravity at 20 percent of the wing mean aerodynamic chord. Only a small amount of data was obtained from the test because, immediately after booster separation at a Mach number of 0.88, the configuration was directionally unstable and diverged in sideslip. Simultaneous with the sideslip divergence, the model became longitudinally unstable at 3 degree angle of attack and -6 degree sideslip and diverged in pitch to a high angle of attack. During the pitch-up the free-floating horizontal tail became unstable at 5 degree angle of attack and the tail drifted against its positive deflection limit.
Preliminary Results Obtained from Flight Test of a Rocket Model Having the Tail Only of the Grumman XF10F Airplane Configuration, TED No. NACA DE 354
A flight test was made to determine the servoplane effectiveness and stability characteristics of the free-floating horizontal stabilizer to be used on the XF10F airplane. The results of this test indicate that servoplane effectiveness is practically constant through the speed range up to a Mach number of 1.15, and the stabilizer static stability is satisfactory. A loss of damping occurs over a narrow Mach number range near M = 1.0, resulting in dynamic instability of the stabilizer in this narrow range. Above M = 1.0 there is a gradual positive trim change of the stabilizer.
Preliminary results of a determination of temperatures of flames by means of K-band microwave attenuation
The temperature effects on the attenuation of K-band microwaves, at a frequency of 26,500 plus or minus 30 megacycles per second, through natural-gas and propane flames containing added alkali halide salts, were investigated over a temperature range from 1900 to 2500 K. The preliminary data of this investigation indicated that the attenuation varies appreciably with the sodium-line-reversal temperatures of the flames and is independent of the particular hydrocarbon fuels that were used for temperature sources and of the particular halide components of the compounds used in the concentrations employed to produce easily measurable attenuation. A reproducibility of plus or minus 25 K was obtainable.
Preliminary Results of a Flight Investigation of 1/6-Scale Rocket-Powered Models of the Bell MX-776 to Determine Aileron Rolling Effectiveness and Total Drag
An experimental investigation of the variation of aileron rolling effectiveness and total drag with Mach number has been made using 1/6-scale rocket-propelled models of the Bell MX-776. Three models having constant-chordwise-thickness full-span aileron at approximate deflections of 2 deg, 5 deg, and 15 deg have been flown. Positive control effectiveness over the Mach number range between approximately 0.5 and 1.2 was obtained from the models and no indication of reversal of effectiveness was encountered. The ratio of tip helix angle to aileron deflection indicated a decrease in proportional rolling effectiveness with increasing deflections in the Mach number range from approximately 0.7 to 1.0. A drag rise of about 125 percent in the transonic region between Mach numbers of 0.85 and 1.02 followed by a gradual decrease at higher speeds was revealed.
Preliminary results of a flight investigation to determine the effect of negative flap deflection on high-speed longitudinal-control characteristics
No Description Available.
Preliminary Results of a Free-Flight Investigation of the Static Stability and Aileron Control Characteristics of 1/6 Scale Models of the Bell MX-776
An investigation of the static longitudinal stability, static directional stability, and aileron control characteristics at transonic and supersonic speeds is being made of 1/6 scale rocket-propelled model of the Bell MX-776. A stability investigation has been made of two symmetrical models with controls undeflected and centers of gravity one-half and one-body diameter, respectively, ahead of the equivalent design center-of-gravity location of the full-scale version. Both models developed large normal-force coefficients in both the subsonic and supersonic ranges which indicated longitudinal instability at low angles of attack. The side-force coefficients were small for both models and indicated that the models were directionally stable. A possible tendency toward dynamic directional instability in the transonic region was indicated by short-period oscillations of the side forces. The results showed a partial-span inboard aileron to be ineffective or to cause negative control in the the transonic region when deflected approximately 5 deg but not when deflected 10 deg. An investigation of drag showed it to increase with a rearward movement of the center of gravity. This indicates an increase in the trim angle of attack as could be caused by a decrease in static stability.
Preliminary Results of a Survey for Thick High-Calcium Limestone Deposits in the United States
From introduction: This report contains the results of a preliminary study of limestone deposits in the United States and Alaska for the purpose of selecting those deposits of sufficient size, relief, and purity in which to conduct an underground nuclear test.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. I - Pressure and Temperature Distributions
Pressures and temperatures throughout the X24C-4B turbojet engine are presented in both tabular and graphical forms to show the effect of altitude, flight Mach number, and engine speed on the internal operation of the engine. These data were obtained in the NACA Cleveland altitude wind tunnel at simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.08, and engine speeds from 4000 to 12,500 rpm. Location and detail drawings of the instrumentation installed at seven survey stations in the engine are shown. Application of generalization factors to pressures and temperatures at each measuring station for the range of altitudes investigated showed that the data did not generalize above an altitude of 25,000 feet. Total-pressure distribution at the compressor outlet varied only with change in engine speed. At altitudes above 35,000 feet and engine speeds above 11,000 rpm, the peak temperature at the turbine-outlet annulus moved inward toward the root of the blade, which is undesirable from blade-stress considerations. The temperature levels at the turbine outlet and the exhaust-nozzle outlet were lowered as the Mach number was increased. The static-pressure measurements obtained at each stator stage of the compressor showed a pressure drop through the inlet guide vanes and the first-stage rotor at high engine speeds. The average values measured by the manufacturer's instrumentation werein close agreement with the average values obtained with NACA instrumentation.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. II - Engine Performance, II, Engine Performance
An investigation was conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of the X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet,simulated flight Mach numbers from 0 to 1.08, and engine speeds from 4000 to 12,500 rpm. Performance data are presented to show graphically the effects of altitude at a flight Mach number of 0.25 and of flight Mach number at an altitude of 25,000 feet. The performance data are generalized to show the applicability of methods used to determine performance at any altitude from data obtained at a given altitude. A complete tabulation of performance data, as well as lubrication- and fuel- system data, is presented.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. IV - Performance of Modified Compressor, Part 4, Performance of Modified Compressor
The performance of the 11-stage axial-flow compressor, modified to improve the compressor-outlet velocity, in a revised X24C-4B turbojet engine is presented and compared with the performance of the compressor in the original engine. Performance data were obtained from an investigation of the revised engine in the MACA Cleveland altitude wind tunnel. Compressor performance data were obtained for engine operation with four exhaust nozzles of different outlet area at simulated altitudes from 15,OOO to 45,000 feet, simulated flight Mach numbers from 0.24 to 1.07, and engine speeds from 4000 to 12,500 rpm. The data cover a range of corrected engine speeds from 4100 to 13,500 rpm, which correspond to compressor Mach numbers from 0.30 to 1.00.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X24C-4B Turbojet Engine. V - Performance of Modified Engine, V, Performance of Modified Engine
An investigation has been conducted in the NACA Cleveland altitude wind tunnel to evaluate the performance characteristics of a modified X24C-4B turbojet engine over a range of simulated altitudes from 5000 to 45,000 feet, simulated flight Mach numbers from 0.25 to 1.07, and engine speeds from 4000 to 12,500 rpm. The engine was modified by the manufacturer to improve the velocity and temperature profiles within the engine. Performance data are graphically presented to show the effect of altitude at a flight Mach number of 0.25 and the effect of flight Mach number at an altitude of 25,000 feet. Original and modified engine performances for several specific operating conditions are compared. A complete tabulation of average pressures and temperatures throughout the engine, performance data, and lubrication and fuel-system data is presented.
Preliminary Results of Altitude-Wind-Tunnel Investigation of X34C-4B Turbojet Engine. III - Compressor Performance, 3, Compressor Performance
The performance of the 11-stage axial-flow compressor in the X24C-4B turbojet engine was analyzed on the basis of results obtained from an investigation of the complete engine in the NACA Cleveland altitude wind tunnel. The engine was operated with four, exhaust nozzles of different outlet area over a range of engine speeds from 6000 to 12,500 rpm, corrected engine speeds from approximately 6100 to 13,600 rpm, and compressor Mach numbers from 0.45 to 1.00. Data are presented for engine operation over a range of simulated altitudes from 15,000 to 45,000 feet and simulated flight Mach numbers from 0.24 to 1.08.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine, 3, Pressure and Temperature Distributions
An altitude-wind-tunnel investigation of a TG-100A gas turbine-propeller engine was performed. Pressure and temperature data were obtained at altitudes from 5000 to 35000 feet, compressor inlet ram-pressure ratios from 1.00 to 1.17, and engine speeds from 800 to 13000 rpm. The effect of engine speed, shaft horsepower, and compressor-inlet ram-pressure ratio on pressure and temperature distribution at each measuring station are presented graphically.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine, 4, Compressor and Turbine Performance Characteristics
As part of an investigation of the performance and operational characteristics of the TG-100A gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100R. The highest compressor pressure ratio was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine II - Windmilling Characteristics
An investigation was conducted to determine the operational and performance characteristics of the TG-100A gas turbine-propeller engine II. Windmilling characteristics were deterined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of a TG-100A Gas Turbine-Propeller Engine, V, Combustion-Chamber Characteristics
An investigation to determine the performance and operational characteristics of the TG-1OOA gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet rm-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and.correcte& horsepower. For the range of corrected engine speeds investigated, over-all total-pressure-loss ratio, cycle efficiency, ana the frac%ional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. The scatter of combustion- efficiency data tended to obscure any effect of altitude or ram-pressure ratio. For the range of corrected horse-powers investigated, the total-pressure-loss ratio an& the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horse-powers investigated at all corrected engine speeds.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 1, Performance Characteristics
A preliminary investigation of an axial-flow gas turbine-propeller engine was conduxted. Performance data were obtained for engine speeds from 8000 to 13,000 rpm and altitudes from 5000 to 35,000 feet and compressor inlet ram pressure ratios from 1.00 to 1.17.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 2, Windmilling Characteristics
Wind tunnel investigations were performed to determine the performance properties of an axial-flow gas turbine-propeller engine II. Windmilling characteristics were determined for a range of altitudes from 5000 to 35,000 feet, true airspeeds from 100 to 273 miles per hour, and propeller blade angles from 4 degrees to 46 degrees.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 3, Pressure and Temperature Distributions
Performance properties and operational characteristics of an axial-flow gas turbine-propeller engine were determined. Data are presented for a range of simulated altitudes from 5,000 to 35,0000 feet, compressor inlet- ram pressure ratios from 1.00 to 1.17, and engine speeds from 8000 to 13,000 rpm.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 4, Compressor and Turbine Performance Characteristics
As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.
Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine, 5, Combustion-Chamber Characterisitcs
An investigation to determine the performance and operational characteristics of an axial-flow gas turbine-propeller engine was conducted in the Cleveland altitude wind tunnel. As part of this investigation, the combustion-chamber performance was determined at pressure altitudes from 5000 to 35,000 feet, compressor-inlet ram-pressure ratios of 1.00 and 1.09, and engine speeds from 8000 to 13,000 rpm. Combustion-chamber performance is presented as a function of corrected engine speed and corrected horsepower. For the range of corrected engine speeds investigated, overall total-pressure-loss ratio, cycle efficiency, and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers were unaffected by a change in altitude or compressor-inlet ram-pressure ratio. For the range of corrected horsepowers investigated, the total-pressure-loss ratio and the fractional loss in cycle efficiency resulting from pressure losses in the combustion chambers decreased with an increase in corrected horsepower at a constant corrected engine speed. The combustion efficiency remained constant for the range of corrected horsepowers investigated at all corrected engine speeds.
Preliminary results of an investigation at transonic speeds to determine the effects of a heated propulsive jet on the drag characteristics of a related series of afterbodies
Preliminary results are presented from an investigation to determine the influence of afterbody geometry on the effects of a sonic propulsive jet at transonic speeds. The results presented are base pressure coefficient and afterbody pressure-drag coefficient as a function of jet pressure ratio for different values of Mach number and jet temperature. Geometric parameters investigated include boattail angle, jet-to-model diameter ratio, and jet-to-base diameter ratio.
Preliminary Results of an Investigation by the Wing-Flow Method of the Longitudinal Stability Characteristics of a 1/50-Scale Semispan Model of the McDonnell XP-88 Airplane
This paper presents the results of measurements of longitudinal stability of a 1/50-scale model of the XP-88 airplane by the wing-flow method. Lift, rolling-moment, hinge-moment, and pitching-moment characteristics as well as the downwash at the tail were measured over a Mach number range from approximately 0.5 to 1.05 at Reynolds numbers below 1,000,000. No measurements of drag were obtained. No abrupt changes due to Mach number were noted in any of the parameters measured. The data indicated that the wing was subject to early tip stalling; that the tail effectiveness decreased gradually with increasing Mach number up to M = 0.9, but increased again at higher Mach numbers; that the variation of downwash with angle of attack did not change appreciably with Mach number except between 0.95 and 1.0 where d(epsilon)/d(alpha), decreased from 0.46 to 0.32; that at zero lift with a stabilizer setting of -1.5 deg there was a gradually increasing nosing-up tendency with increasing Mach number; and that the control-fixed stability in maneuvers at constant speed gradually increased with increasing Mach number.
Preliminary results of an investigation of the effects of spinner shape on the characteristics of an NACA D-type cowl behind a three-blade propeller, including the characteristics of the propeller at negative thrust
No Description Available.
Preliminary Results of British Nene II Engine Altitude-Chamber Performance Investigation. I - Altitude Performance Using Standard 18.75-Inch-Diameter Jet Nozzle, 1, Altitude Performance Using Standard 18.75-Inch-Diameter Jet Nozzle
An investigation is being conducted to determine the altitude performance characteristics of the British Nene II engine and its components. The present paper presents the preliminary results obtained using a standard jet nozzle. The test results presented are for conditions simulating altitudes from sea level to 60,000 feet and ram pressure ratios from 1.0 to 2.3. These ram pressure ratios correspond to flight Mach numbers between zero and 1.16 assuming a 100 percent ram recovery.
Preliminary Results of Cyclical De-Icing of a Gas-Heated Airfoil
An NACA 65(sub 1)-212 airfoil of 8-foot chord was provided with a gas-heated leading edge for investigations of cyclical de-icing. De-icing was accomplished with intermittent heating of airfoil segments that supplied hot gas to chordwise passages in a double-skin construction. Ice removal was facilitated by a spanwise leading-edge parting strip which was continuously heated from the gas-supply duct. Preliminary results demonstrate that satisfactory cyclical ice removal occurs with ratios of cycle time to heat-on period (cycle ratio) from 10 to 26. For minimum runback, efficient ice removal, and minimum total heat input, short heat-on periods of about 15 seconds with heat-off periods of 260 seconds gave the best results. In the range of conditions investigated, the prime variables in the determination of the required heat input for cyclical ice removal were the air temperature and the cycle ratio; heat-off period, liquid water content, airspeed, and angle of attack had only secondary effects on heat input rate.
Preliminary Results of Heat Transfer from a Stationary and Rotating Ellipsoidal Spinner
Convective heat-transfer coefficients in dry air were obtained for an ellipsoidal spinner of 30-inch maximum diameter for both stationary and rotating operation over a range of conditions including airspeeds up to 275 miles per hour, rotational speeds up to 1200 rpm, and angles of attack of zero and 40 The results are presented in terms of Nusselt numbers, Reynolds numbers, and convective heat-transfer coefficients. The studies included both uniform heating densities over the spinner and uniform surface temperatures.. In general, the results showed that rotation will increase the convective heat transfer from a spinner, especially in the turbulent-flow regions. Rotation of the spinner at 1200 rpm and at a free-stream velocity of 275 miles per hour increased the Nusselt number parameter in the turbulent-flow region by 32 percent over that obtained with a stationary spinner; whereas in the nose region, where the flow was laminar, an increase of only 18 percent was observed. Transition from laminar to turbulent flow occurred over a large range of Reynolds numbers primarily because of surface roughness of the spinner. Operation at an angle of attack of 40 had only small effects on the local convective heat transfer for the model studied.
Preliminary results of horizontal-tail load measurements of the Bell X-5 research airplane
No Description Available.
Preliminary results of NACA transonic flights of the XS-1 airplane with a 10-percent-thick wing and 8-percent-thick horizontal tail
No Description Available.