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High-Speed Load Distribution of the Wing of a 3/16-Scale Model of the Douglas XSB2D-1 Airplane with Flaps Deflected
"The tests reported herein were made for the purpose of determining the high-speed load distribution on the wing of a 3/16 scale model of the Douglas XSB2D-1 airplane. Comparisons are made between the root bending moment and section torsional moment coefficients as obtained experimentally and derived analytically. The results show good correlation for the bending moment coefficients but considerable disagreement for the torsional moment coefficients, the measured moments being greater than the analytical moments. The effects of Mach number on both the bending moment and torsional moment coefficients were small" (p. 1).
Tank Tests of a 1/7-Size Dynamic Model of the Grumman XJR2F-1 Amphibian to Determine the Effect of Slotted- and Split-Type Flaps on Take-Off Stability - NACA Model 212, TED No. NACA 2378
From Summary: "Additional tests of a 1/7-size model of the Grumman XJR2F-1 amphibian were made in Langley tank no. 1 to compare the behavior during take-off of the model equipped with split- and slotted-type flaps. The slotted flag had a large effect on locating the forward center-of-gravity limits for stable take-offs. Stable take-offs within the normal operating range of positions of the center of gravity could be made with the split flaps deflected 45 degrees or with the slotted flaps deflected less than 20 degrees."
Cyclic Engine Test of Cast Vitallium Turbine Buckets - I
"An investigation was conducted to correlate the engine service performance of cast Vitallium turbine buckets with standard laboratory metallurgical data. Data were obtained from four I-40 turbine wheels of Timken alloy with cast Vitallium buckets. In order to accelerate bucket deterioration, the turbine wheels were subjected to 20-minute cycles consisting of 5 minutes at idle and 15 minutes at rated speed" (p. 1).
Analytical Comparison of a Standard Turbojet Engine, a Turbojet Engine with a Tail-Pipe Burner, and a Ram-Jet Engine
From Introduction: "Experimental investigations (reference 1) have shown that in some cases the thrust can be more than doubled by means of tail-pipe burning. A comparison is made of a standard turbojet engine, whose thrust is augmented by tail-pipe burning, and a ram-jet engine. The performance characteristics for the ram-jet engine were computed entirely from theoretical considerations and on the assumption that the burner-inlet velocity was constant."
Two-Dimensional Wind-Tunnel Investigation of Modified NACA 65(sub 112)-111 Airfoil with 35-Percent-Chord Slotted Flap to Determine Pitching-Moment Characteristics and Effects of Roughness
From Summary: "An investigation has been made in the Langley two-dimensional low-turbulence pressure tunnel to develop the optimum configuration of a 0.35-chord slotted flap on an NACA 65(sub 1120)-111 airfoil section modified by removing the trailing-edge cusp. The section pitching-moment characteristics and the effects of standard roughness on the section characteristics were determined for the flap retracted at Reynolds numbers ranging from 3.0 x 10(exp 6) to 9.0 x 10(exp 6)."
Computed Temperature Distribution and Cooling of Solid Gas-Turbine Blades
"Computations were made to determine the temperature distribution and cooling of solid gas-turbine blades.A range of temperatures was used from 1500 degrees to 2500 degrees F, blade-root temperatures from 100 degrees to 1000 degrees F, blade thermal conductivity from 8 to 220 BTU/(hr)(sq ft)(degrees F/ft), and net gas to metal heat transfer coefficients from 75 to 250 BTU/(hr)(sq ft)(degrees F)" (p. 1).
Cooling of Gas Turbines 1 - 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
"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" (p. 1).
Cooling of Gas Turbines, 3, Analysis of Rotor and Blade Temperatures in Liquid-Cooled Gas Turbines
A theoretical analysis of the radial temperature distribution through the rotor and constant cross sectional area blades near the coolant passages of liquid cooled gas turbines was made. The analysis was applied to obtain the rotor and blade temperatures of a specific turbine using a gas flow of 55 pounds per second, a coolant flow of 6.42 pounds per second, and an average coolant temperature of 200 degrees F. The effect of using kerosene, water, and ethylene glycol was determined. The effect of varying blade length and coolant passage lengths with water as the coolant was also determined. The effective gas temperature was varied from 2000 degrees to 5000 degrees F in each investigation.
Development of Inboard Nacelle for the XB-36 Airplane
From Summary: "A series of investigations of several 1/14-scale models of an inboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels. The purpose of these investigations was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. As a result of these investigations, a nacelle was developed which had external drag coefficients considerably lower than the original basic form with the external nacelle drag approximately one-half to two-thirds of those of conventional tractor designs. The largest reductions in drag resulted from sealing the gaps between the wing flaps and nacelle, reducing the thickness of the nacelle training-edge lip, and bringing the under-wing air inlet to the wing leading edge."
Investigation of Pressure Losses in Several Turbosupercharger Nozzle Hoses
Memorandum presenting surveys of the impact pressure of the flow to obtain information for determining pressure losses in four different turbosupercharger nozzle boxes. The data indicated substantial differences in total head loss among the boxes and the existence of sharply defined high-loss regions in portions of the nozzle annulus.
Theoretical Evaluation of Methods of Cooling the Blades of Gas Turbines
A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades. The results indicated that cooling of the root of the blade, shortening the blade, and cooling hollow blades internally with air or liquid offer possibilities of substantial increases in permissible gas temperatures.
Full-Scale Investigation of the Maximum Lift and Flow Characteristics of an Airplane Having Approximately Triangular Plan Form
Report discussing an investigation of the DM-1 glider, which has an approximately triangular plan form, with auxiliary studies of a model of triangular wings. The pitching-moment coefficient, drag coefficient, and angle of attack with the lift coefficient are provided. Results indicated that the angles of descent without power are likely to be prohibitive and airplanes with the tested type of wings will not be able to land safely without power.
Investigation of Rim Cracking in Turbine Wheels with Welded Blades
Rim cracking in turbine wheels with welded blades was evaluated. The problem is explained on the basis of the occurrence of plastic flow in the rim during transient starting conditions when thermal compressive stresses resulting from high-temperature gradients exceed the proportional elastic limit of the material.
Aerodynamic Characteristics at High Speeds of Full-Scale Propellers Having Different Shank Designs
"Tests of two 10-foot-diameter two-blade propellers which differed only in shank design have been made in the Langley 16-foot high-speed tunnel. The propellers are designated by their blade design numbers, NACA 10-(5)(08)-03, which had aerodynamically efficient airfoil shank sections, and NACA 10-(5)(08)-03R which had thick cylindrical shank sections typical of conventional blades. The propellers mere tested on a 2000-horsepower dynamometer through a range of blade-angles from 20 degrees to 55 degrees at various rotational speeds and at airspeeds up to 496 miles per hour" (p. 1).
Comparison of Wind-Tunnel Predictions with Flight Measurements of the Longitudinal-Stability and -Control Characteristics of a Douglas BTD-1 Airplane
"Low Mach number longitudinal-stability and control characteristics as predicted by use of wind tunnel data from a powered 3/16-scale model are compared with flight test measurements of a Navy BTD-1 airplane. The accuracy of the wind tunnel data and the discrepancies involved in attempting to correlate with flight data are discussed and analyzed. The comparison showed that wind tunnel predictions were, in general, in good agreement with flight test data" (p. 1).
Flight Comparison of Performance and Cooling Characteristics of Exhaust-Ejector Installation with Exhaust-Collector-Ring Installation
Flight and ground investigations have been made to compare an exhaust-ejector installation with a standard exhaust-collector-ring installation on air-cooled aircraft engines in a twin-engine airplane. The ground investigation allowed that, whereas the standard engine would have overheated above 600 horsepower, the engine with exhaust ejectors cooled at take-off operating conditions at zero ram. The exhaust ejectors provided as much cooling with cowl flaps closed as the conventional cowl flaps induced when full open at low airspeeds. The propulsive thrust of the exhaust-ejector installation was calculated to be slightly less than the thrust of the collector-ring-installation.
High-Speed Wind-Tunnel Tests of a Model of the Lockheed YP-80A Airplane Including Correlation with Flight Tests and Tests of Dive-Recovery Flaps
"This report contains the results of tests of a 1/3-scale model of the Lockheed YP-90A "Shooting Star" airplane and a comparison of drag, maximum lift coefficient, and elevator angle required for level flight as measured in the wind tunnel and in flight. Included in the report are the general aerodynamic characteristics of the model and of two types of dive-recovery flaps, one at several positions along the chord on the lower surface of the wing and the other on the lower surface of the fuselage. The results show good agreement between the flight and wind-tunnel measurements at all Mach numbers" (p. 1).
An Investigation of the Effects of Sweep on the Characteristics of a High-Aspect-Ratio Wing in the Langley 8-Foot High-Speed Tunnel
"An untwisted wing, which when unswept has an NACA 65-210 section, an aspect ratio of 9.0 and a taper ration of 2.5:1.0, has been tested with no sweep, and 30 deg and 45 deg of sweepback and sweepforward in conjunction with a typical fuselage at Mach numbers from 0.60 to 0.96 at angles of attack generally between -2 deg and 10 deg in the Langley 8-foot high-speed tunnel. Sweep was obtained by rotating the wing semispans about a point in the plane of symmetry. The normal-force, pitching-moment, profile-drag, and loading characteristics for the wings have been obtained from pressure measurements and wake surveys" (p. 1).
An Analytical Investigation of the Heat Losses from a U.S. Navy K-Type Airship
From Summary: "The heat losses from the envelope surface of a U.S. Navy K-type airship are evaluated to determine if the use of heat is a feasible means of preventing ice and snow accumulations on lighter-than-air craft during flight and when moored uncovered. Consideration is given to heat losses in clear air (no liquid water present in the atmosphere) and in probable conditions of icing and snow. The results of the analysis indicate that the amount of heat required in flight to raise the surface temperature of the entire envelope to the extent considered adequate for ice protection, based on experience with tests of heavier-than-air craft, is very large."
Tank Tests of 1/5.5-Scale Forward Dynamic Model of the Columbia XJL-1 Amphibian - Langley Tank Model 208, TED No. NACA 2336
Tests of a powered dynamic model of the Columbia XJL-1 amphibian were made in Langley tank no.1 to determine the hydrodynamic stability and spray characteristics of the basic hull and to investigate the effects of modifications on these characteristics. Modifications to the forebody chime flare, the step, and the afterbody, and an increase in the angle of incidence of the wing were included in the test program. The seaworthiness and spray characteristics were studied from simulated taxi runs in smooth and rough water. The trim limits of stability, the range of stable positions of the enter of gravity for take-off, and the landing stability were determined in smooth water. The aerodynamic lift, pitching moment, and thrust were determined at speeds up to take-off speed.
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
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.
Calculations of the Supersonic Wave Drag of Nonlifting Wings with Arbitrary Sweepback and Aspect Ratio: Wings Swept Behind the Mach Lines
"On the basis of a recently developed theory for finite sweptback wings at supersonic speeds, calculations of the supersonic wave drag at zero lift were made for a series of wings having thin symmetrical biconvex sections with untapered plan forms and various angles of sweepback and aspect ratios. The results are presented in a unified form so that a single chart permits the direct determination of the wave drag for this family of airfoils for an extensive range of aspect ratio and sweepback angle for stream Mach numbers up to a value corresponding to that at which the Mach line coincides with the wing leading edge. The calculations showed that in general the wave-drag coefficient decreased with increasing sweepback" (p. 1).
Investigation of High-Performance Fuels in Multicylinder and in Single-Cylinder Engines at High and Cruising Engine Speeds
"An investigation was conducted to compare the knock-limited performance of a 20-percent triptane blend in 28-K fuel with that of 28-R and 33-R fuels at high engine speeds, cruising speeds, and two compression ratios in an K-1830-94 multicylinder engine. Data were obtained with the standard compression ratio of 6.7 and with a compression ratio of 3.0. The three fuels were investigated at engine speeds of 1800, 2250, 2600, and 2800 rpm at high and low blower ratios" (p. 1).
Preliminary Tests at Supersonic Speeds of Triangular and Swept-Back Wings
Report presenting testing of a series of thin, triangular plan-form wings, including eight triangular wings of vertex angles and three swept-back wings with circular-arc sections. Results regarding lift, center of pressure, and ideal operation of different types of wings are provided.
Drag characteristics of rectangular and swept-back NACA 65-009 airfoils having aspect ratios of 1.5 and 2.7 as determined by flight tests at supersonic speeds
Report presenting tests to determine the effects of sweepback angle and aspect ratio on the drag of an NACA 65-009 airfoil at supersonic speeds. The results indicated that for the range of Mach numbers investigated, increasing the sweepback angle and decreasing the aspect ratio reduced the value of the wing drag coefficient.
Minimum Specific Fuel Consumption of a Liquid-Cooled Multicylinder Aircraft Engine as Affected by Compression Ratio and Engine Operating Conditions
From Summary: "An investigation was conducted on a 12-cylinder V-type liquid-cooled aircraft engine of 1710-cubic-inch displacement to determine the minimum specific fuel consumption at constant cruising engine speed and compression ratios of 6.65, 7.93, and 9.68. At each compression ratio, the effect.of the following variables was investigated at manifold pressures of 28, 34, 40, and 50 inches of mercury absolute: temperature of the inlet-air to the auxiliary-stage supercharger, fuel-air ratio, and spark advance. Standard sea-level atmospheric pressure was maintained at the auxiliary-stage supercharger inlet and the exhaust pressure was atmospheric."
The effect of high solidity on propeller characteristics at high forward speeds from wind-tunnel tests of the NACA 4-(3)(06.3)-06 and NACA 4-(3)(06.4)-09 two-blade propellers
From Summary: "Tests of two-blade propellers having the NACA 4-(3)(06.3)-06 and NACA 4-(3)(06.4)-09 blade designs (blade activity factors of 179 and 263, respectively) have been made in the Langley 8-foot high-speed tunnel through a range of blade angle from 20 degrees to 70 degrees for free-stream Mach numbers from 0.165 to 0.725 to determine the effects of high solidity and compressibility on propeller characteristics. The tests are part of a general investigation of propellers at high forward speeds. Results previously reported for similar tests of two-blade propellers having the NACA 4-308-03 and NACA 4-308-045 blade designs (blade activity factors of 87 and 133, respectively) are included for comparison. The results showed that the 0.06- and 0.09-solidity blades, although producing efficiencies of the order of 90 percent, were less efficient than blades of conventional solidity."
Lateral Stability Characteristics of a 1/8.33-Scale Powered Model of the Republic XF-12 Airplane
"The XF-12 airplane is a high-performance photo-reconnaissance aircraft designed for the Army Air Forces by the Republic Aviation Corporation. An investigation of a 1/8.33 - scale powered model was made in the Langley l9-foot pressure tunnel to obtain information relative to the aerodynamic design of the airplane. The model was tested with and without the original vertical tail. and with two revised tails. For the revised tail no. 1, the span of the original vertical .tail was increased about 15 percent and the portion of the vertical tail between the stabilizer and fuselage behind the rudder hinge line was allowed to deflect simultaneously with the main rudder" (p. 1).
Determination of Ram-Jet Combustion-Chamber Temperatures by Means of Total-Pressure Surveys
"A method is described by which the total temperature of the gases at the combustion-chamber outlet of a ram-jet engine may be determined from the loss in total pressure measured across the combustion chamber. A working chart is presented by means of which the ratio of the total temperature of the gases at the combustion-chamber outlet to the total temperature of the gases at the combustion-chamber inlet may be determined from the measured loss of total pressure across the combustion chamber and the known values of air flow, total pressure, and total temperature at the combustion-chamber inlet. Values of total-temperature ratio across the combustion chamber of a 20-inch ram jet were obtained in the Cleveland altitude wind tunnel over a range of pressure altitudes from 6000 to 15,000 feet" (p. 1).
Summary of Available Data Relating to Reynolds Number Effects on the Maximum Lift Coefficients of Swept-Back Wings
The available foreign and American data relating to Reynolds number effects on the maximum lift coefficients of swept-back wings are summarized and discussed. The data show that at low Reynolds numbers (below about 2.0 x 10(exp 6)) higher maximum lift coefficients were measured in most cases for moderately swept-back wings than for unswept wings of similar plan form; at high Reynolds numbers, however, increasing sweepback resulted in decreasing maximum lift coefficients. A smaller rate of increase of the maximum lift coefficient with Reynolds number was measured for the swept-back wings than for similar unswept wings in the critical range of Reynolds number. Increasing the Reynolds number resulted in decreases in the maximum lift coefficients of the two wings of approximately triangular plan form that were investigated.
Investigation of the Trim Characteristics of a 1/20-Scale Model of the Fleetwings XBTK-1 Airplane over a Wide Range of Angles of Attack
"Tests of a 1/20-scale model of the Fleetwings XBTK-1 airplane have been performed in the Langley 15-foot free-spinning tunnel to determine the trim tendencies of the airplane at attitudes above the stall. The results of the tests indicated that the model would trim longitudinally only in the normal range of angles of attack and that the yaw trim tendencies for such longitudinal trim conditions were normal. Although wide oscillations in yaw were noted for some conditions, they occurred at angles of attack larger than those indicated as possible for longitudinal trim and spin equilibrium" (p. 1).
Comparative Drag Measurements at Transonic Speeds of an NACA 65-006 Airfoil and a Symmetrical Circular-Arc Airfoil
Report presenting measurements made at transonic speeds by the freely-falling-body method to compare the drag of a rectangular plan-form airfoil of aspect ratio 7.6 with an NACA 65-006 airfoil section. Results regarding the velocity measurements, base-pressure measurements, and airfoil drag measurements are provided.
Ditching Tests with a 1/12-Scale Model of the Army A-26 Airplane in Langley Tank No. 2 and on an Outdoor Catapult
Tests were conducted in calm water in Langley tank no. 2 and in calm and rough water at an outdoor catapult in order to determine the best way to make a forced landing of an Army A-26 airplane and to determine its probable ditching behavior. These tests were requested by the Air Materiel Command, Army Air Forces, in their letter of March 26, 1943, WEL:AW:50.
Free-Spinning and Tumbling Tests of a 1/16-Scale Model of the McDonnell XP-85 Airplane
The teat results showed that with either of the three tail arrangements, the model usually spun in flat attitudes with oscillations about the lateral and longitudinal axes. In general, full reversal of the rudder pedals did not stop the spinning rotation. To make the model satisfactorily meet-the spin-recovery requirements it was found that installation of either a very large ventral fin (l7.9 square feet, full scale) below the tail or a somewhat smaller ventral fin and rudder (12.4 square feet, total . full-scale area) with a rudder throw of at least +/-22deg was required. Either a 21.3-foot tail parachute or a 6.4-foot wing-tip parachute (drag coefficient approximately 0.70) appears necessary as an emergency spin-recovery device during demonstration spins.
Drag Measurements of Symmetrical Circular-Arc and NACA 65-009 Rectangular Airfoils Having an Aspect Ratio of 2.7 as Determined by Flight Tests at Supersonic Speeds
Report discussing testing to determine the drag characteristics at zero lift of a wing with a circular-arc airfoil section with a maximum thickness of 9 percent chord. The results were compared to previous testing on an NACA 65-009 airfoil. It was found that the NACA airfoil had lower drag coefficients than the circular-arc airfoil tested in this experiment.
Ditching Tests of a 1/18-Scale Model of the Navy XP4M-1 Airplane in Langley Tank No. 2 and on an Outdoor Catapult, TED No. NACA 2362
From Summary: "Tests with a dynamically similar model of the Navy XP4M-1 airplane were made to determine the best way to land the airplane in calm and rough water, to determine its probable ditching performance, and to determine practicable modifications which could be incorporated in the design of the airplane that would improve its ditching characteristics. The results were obtained by making visual observations, by recording longitudinal decelerations ,and by taking motion pictures of the landings. A list of conclusions from the test results is included."
Evaluation of Gust and Draft Velocities from Flights of P-61C Airplanes within Thunderstorms July 22, 1946 to July 23, 1946 at Orlando, Florida
"The results obtained from measurements of gust and draft velocities within thunderstorms for the period July 22, 1946 to July 23, 1946 at Orlando, Florida, are presented herein. These data are summarized in tables I and II, respectively, and are of the type presented in reference 1 for previous flights. Inspection of photo-observer records for the flights indicated that no data on ambient air temperature variations within thunderstorms were obtained" (p. 1).
Aerodynamic Measurements Made During Navy Investigation of Human Tolerance to Wind Blasts
From Summary: "This report presents the aerodynamic measurements made during a Navy investigation conducted in the Langley 8-foot high speed tunnel to determine the actual human tolerance to wind blasts."
Knock-Limited Performance of Triptane and 28-R Fuel Blends as Affected by Changes in Compression Ratio and in Engine Operating Variables
From Summary: "A knock-limited performance investigation was conducted on blends of triptane and 28-P fuel with a 12-cylinder, V-type, liquid-cooled aircraft engine of 1710-cubic-inch displacement at three compression ratios: 6.65, 7.93, and 9.68. At each compression ratio, the effect of changes in temperature of the inlet air to the auxiliary-stage supercharger and in fuel-air ratio were investigated at engine speeds of 2280 and. 3000 rpm. The results show that knock-limited engine performance, as improved by the use of triptane, allowed operation at both take-off and cruising power at a compression ratio of 9.68."
Wind-Tunnel Investigation of Wing Inlets for a Four-Engine Airplane
Report presenting an investigation in the propeller-research tunnel to develop wing-leading-edge inlets for locations between the inboard and outboard nacelles on each wing of a four-engine airplane. Testing was performed on the basic wing and original inlet as well as NACA-developed inlets for two versions of the airplane.
Evaluation of Gust and Draft Velocities from Flights of P-61C Airplanes within Thunderstorms September 11, 1946 to September 16, 1946 at Orlando, Florida
"The results obtained from measurements of gust velocities, draft velocities, and ambient-air temperature within thunderstorms for the period from September 11, 1946 to September 16, 1946 at Orlando, Florida are presented herein. These data are summarized in.and presented" (p. 1).
Two-dimensional wind-tunnel investigation at high Reynolds numbers of two symmetrical circular-arc airfoil sections with high-lift devices
Report presenting an investigation of two symmetrical circular-arc airfoils of 6 and 10 percent thickness equipped with leading-edge and trailing-edge high-lift devices. A trialing-edge flap, drooped-nose flap, and leading-edge extensible flap were all tested. Results regarding the plain airfoils, airfoils with high-lift devices, combined deflection of high-lift devices, and low-drag-control flaps are provided.
Aerodynamic Characteristics of Three Deep-Stepped Planing-Tail Flying-Boat Hulls
"An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of three deep-stepped planing-tail flying-boat hulls differing only in the amount of step fairing. The hulls were derived by increasing the unfaired step depth of a planing-tail hull of a previous aerodynamic investigation to a depth about 92 percent of the hull beam. Tests were also made on a transverse-stepped hull with an extended afterbody for the purpose of comparison and in order to extend and verify the results of a previous investigation" (p. 1).
Drag of a Wing-Body Configuration Consisting of a Swept-Forward Tapered Wing Mounted on a Body of Fineness Ratio 12 Measured During Free Fall at Transonic Speeds
Report discussing an investigation to determine the drag of a configuration with a body of fitness ratio 12 with stabilizing tail surfaces and a 12-percent-thick 30-degree swept-forward wing using the free-fall method. The drag oft he wing and the total drag were measured separately and compared. The swept-forward wing was found to greatly increase the effect of drag on the body-tail combination.
Ditching Tests with a 1/16-Size Model of the Navy XP2V-1 Airplane at the Langley Tank No. 2 Monorail
"Tests were made with a 1/16 size dynamically similar model of the Navy XP2V-1 airplane to study its performance when ditched. The model was ditched in calm water at the Langley tank no. 2 monorail. Various landing attitudes, speeds, and conditions of damage were simulated" (p. 1).
Evaluation of Gust and Draft Velocities from Flights of P-61C Airplanes Within Thunderstorms September 17, 1946 to September 18, 1946 at Orlando, Florida
The results obtained from measurements of gust velocities, draft velocities, and ambient-air temperature within thunderstorms for the period September 17, 1946 to September 18, 1946 at Orlando, Fla. are presented herein. These data are summarized in tables I, II, and III, respectively, and are of the type presented in reference 1 for previous flights.
Cooling of Gas Turbines, 2, Effectiveness of Rim Cooling of Blades
An analysis is presented of rim cooling of gas-turbine blades; that is, reducing the temperature at the base of the blade (wheel rim), which cools the blade by conduction alone. Formulas for temperature and stress distributions along the blade are derived and, by the use of experimental stress-rupture data for a typical blade alloy, a relation is established between blade life (time for rupture), operating speed, and amount of rim cooling for several gas temperatures. The effect of blade parameter combining the effects of blade dimensions, blade thermal conductivity, and heat-transfer coefficient is determined. The effect of radiation on the results is approximated. The gas temperatures ranged from 1300F to 1900F and the rim temperature, from 0F to 1000F below the gas temperature. This report is concerned only with blades of uniform cross section, but the conclusions drawn are generally applicable to most modern turbine blades. For a typical rim-cooled blade, gas temperature increases are limited to about 200F for 500F of cooling of the blade base below gas temperature, and additional cooling brings progressively smaller increases. In order to obtain large increases in thermal conductivity or very large decreases in heat-transfer coefficient or blade length or necessary. The increases in gas temperature allowable with rim cooling are particularly small for turbines of large dimensions and high specific mass flows. For a given effective gas temperature, substantial increases in blade life, however, are possible with relatively small amounts of rim cooling.
A Preliminary Theoretical Study of Aerodynamic Instability of a Two-Blade Helicopter Rotor
Report presenting numerical results of flutter calculations for a two-blade helicopter using a see-saw-type rotor. The stability condition for oscillatory motion is expressed in terms of a small number of composite parameters that are evaluated from moments of inertia, angle settings, and aerodynamic parameters of a blade. The analysis indicated that a see-saw rotor with a coning angle is more unstable than an airplane wing with the same parameters.
Evaluation of Gust and Draft Velocities from Flights of P-61C Airplanes Within Thunderstorms August 14, 1946 to August 15, 1946 at Orlando, Florida
"Tables I and II of the present paper summarize the gust and draft velocity data for thunderstorm-flights 21 and 22 of August 14, 1946 and August 15, 1946, respectively. These data were evaluated from records of NACA airspeed-altitude and acceleration recorders installed in P-61C airplanes and are of the type presented for previous flights. Table III summarizes the readings of a milliammeter which was used in conjunction with other equipment to indicate ambient-air temperature during thunderstorm surveys. These data were read from photo-observer records and include all cases in which variations of the instrument indications were noted for the present flights" (p. 1).
Force Tests of a 1/5-Scale Model of the McDonnell XP-85 Airplane with Conventional Tail Assembly in the Langley Free-Flight Tunnel
"At the request of the Air Materiel Command, Army Air Forces an investigation of the low-speed, power-off stability and control characteristics of the McDonnell XP-85 airplane is being conducted in the Langley free-flight tunnel. The XP-85 airplane is a parasite fighter carried in a bomb bay of the B-36 airplane. As a part of the investigation a few force tests were made of a 1/5 scale model of the XP-85 with a conventional tail assembly installed in place of the original design five-unit tail assembly" (p. 1).
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