From Summary: "The tumbling characteristics of a 1/20-scale model of the Northrop N-9M airplane have been determined in the Langley 20-foot free-spinning tunnel for various configurations and loading conditions of the model. The investigation included tests to determine whether recovery from a tumble could be effected by the use of parachutes. An estimation of the forces due to acceleration acting on the pilot during a tumble was made. The tests were performed at an equivalent test altitude of 15,000 feet."
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.
Powered models of three different flying boats were landed in oncoming wave of various heights and lengths. The resulting motions and acceleration were recorded to survey the effects of varying the trim at landing, the deceleration after landing, and the size of the waves. One of the models had an unusually long afterbody. The data for landing with normal rates of deceleration indicated that the most severe motions and accelerations were likely to occur at some period of the landing run subsequent to the initial impact.
This paper makes the following assumptions: 1) The flowing gases are assumed to have uniform energy distribution. ("Isoenergetic gas flows," that is valid with the same constants for the the energy equation entire flow.) This is correct, for example, for gas flows issuing from a region of constant pressure, density, temperature, end velocity. This property is not destroyed by compression shocks because of the universal validity of the energy law. 2) The gas behaves adiabatically, not during the compression shock itself but both before and after the shock. However, the adiabatic equation (p/rho(sup kappa) = C) is not valid for the entire gas flow with the same constant C but rather with an appropriate individual constant for each portion of the gas. For steady flows, this means that the constant C of the adiabatic equation is a function of the stream function. Consequently, a gas that has been flowing "isentropically",that is, with the same constant C of the adiabatic equation throughout (for example, in origination from a region of constant density, temperature, and velocity) no longer remains isentropic after a compression shock if the compression shock is not extremely simple (wedge shaped in a two-dimensional flow or cone shaped in a rotationally symmetrical flow). The solution of nonisentropic flows is therefore an urgent necessity.
Flight tests were made of six noninstrumented rocket-powered "Tin Can" models of AAF Project MX-800. Velocity and drag data were obtained by use of CU Doppler radar. The existence of stability and adequate structural strength for flight near zero lift was checked by visual and photographic observation. Drag data obtained during the tests agreed reasonably well with estimates based on experimental data from NACA RM-2 rocket-powered drag research models.
From Summary: "Estimates of the static stick-fixed stability and control characteristics of the Consolidated Vultee model 240 airplane are presented in this report. The estimates are based on tests of a 0.092-scale powered model in the 10-foot wind tunnel of the Guggenheim Aeronautical Laboratory of the California Institute of Technology. Results of the analysis are evaluated in terms of the Army specifications for stability and control characteristics which are more specific and, in general, equal to or more rigid than the Civil Aeronautics Administration requirements."
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.
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.
Tests of a PB2Y-3 flying boat were made at the U.S. Naval Air Station, Patuxent River, Md., to determine its hydrodynamic trim limits of stability. Corresponding tests were also made of a 1/8-size powered dynamic model of the same flying boat in Langley tank no. 1. During the tank tests, the full-size testing procedure was reproduced as closely as possible in order to obtain data for a direct correlation of the results. As a nominal gross load of 66,000 pounds, the lower trim limits of the full-size and model were in good agreement above a speed of 80 feet per second. As the speed decreased below 80 feet per second, the difference between the model trim limits and full-scale trim limits gradually became larger. The upper trim limit of the model with flaps deflected 0 deg was higher than that of the full-size, but the difference was small over the speed range compared. At flap deflections greater than 0 deg, it was not possible to trim either the model of the airplane to the upper limit with the center of gravity at 28 percent of the mean aerodynamic chord. The decrease in the lower trim limits with increase in flap deflection showed good agreement for the airplane and model. The lower trim limits obtained at different gross loads for the full-size airplane were reduced to approximately a single curve by plotting trim against the square root of C(sub delta (sub o)) divided by C(sub V).
From Summary: "Wind-tunnel tests were made of a 1/25 scale model of the Martin JRM-1 airplane to determine: (1) The longitudinal stability and control characteristics of the JRM-1 model near the water and lateral and directional stability characteristics with power while moving on the surface of the water, the latter being useful for the design of tip floats; (2) The stability and stalling characteristics of the wing with a modified airfoil contour; (3) Stability characteristics of a hull of larger design gross weight; The test results indicated that the elevator was powerful enough to trim the original model in a landing configuration at any lift coefficient within the specified range of centers of gravity."
Sinusoidal waves excited by random disturbances have been observed in the laminar boundary layer of a rotating disk at Reynolds numbers in a narrow range below the Reynolds number of transition. Their frequency was found to be approximately a constant times the velocity of the disk. The velocity of the waves at the radius of 11 inches was determined and found to be 0.2 linear velocity of the disk in a downstream direction, making an angle of approximately 14 degrees with the outward drawn radius vector.
Report presenting an investigation of aileron flutter associated with high-speed flight. When the aileron control system was modified by installing a hydraulic irreversible unit, it was possible to delay aileron flutter and reduce its amplitude. Results regarding the aileron flutter with and without the modifications, analysis of the time histories for the control systems, and relation between aileron upfloat and flutter are provided.
Report presenting a wind-tunnel investigation to determine the practicability of the dropped-aileron-type lateral-control device on NACA low-drag airfoils. Section aerodynamic characteristics of an NACA 66(215)-216 airfoil with an aileron of normal profile and an aileron of straight-sided profile with a modified nose shape are presented for various aileron locations, hinge centers, and aerodynamic balances.
From Summary: "Previous tests of blower-blade sections have been extended by a series of tests at 52.5 degrees stagger. The results of these tests have been combined with the earlier test results and are presented in new blade design charts which supersede those previously presented. An investigation in a test blower over a range of stagger from 44 degrees to 65 degrees has shown that for blades at a solidity of 1.0, the two-dimensional cascade data predict the turning angle to within 1/2 degrees."
Report presenting calculations based on a theoretical analysis for a composite engine consisting of a uniflow two-stroke-cycle spark-ignition engine, a compressor, a blowdown turbine, and a steady-flow turbine. Operation of the engine is considered for four cases of gas mixtures and steady-flow turbine temperatures.
Report presenting an investigation of a 42 degree sweptback wing of aspect ratio 4, taper ratio 0.625, and NACA 64-series airfoil sections to study several proposed devices for increasing the maximum lift coefficient and improving the longitudinal stability characteristics of sweptback wings at the stall. Some of the devices tested included leading-edge flaps and slats, trailing-edge split and extended split flaps, upper-surface split flaps, and upper-surface fences. The results regarding the characteristics of the devices and the wing-fuselage combinations are described.
Report presenting a metallurgical examination of the material used in the fabrication of a thermal ice-prevention system after 225 hours of actual flight operation of the heating installation. Only minor corrosion was noted and no impairment of tensile strength was observed.
Report presenting an analysis of the spanwise loading using two different methods on the wing of an airplane for which pressure-distribution measurements were available from flight tests up to a Mach number of 0.866. A comparison between measured and calculated distributions was made on the basis of equal wing-panel normal-force coefficients.
A study of the relations existing among pin-point autoignition, homogeneous autoignition, and knock has been made by means of the NACA high-speed camera and the full-view combustion apparatus. High-speed photographic records of combustion, together with corresponding pressure-time traces, of benzene, 2,2,3-trimethylbutane, S-4, and M-4 fuels at various engine conditions have shown the engine conditions under which each of these phenomena occur and the relation of these phenomena to one another.
Report presenting an investigation of the mechanism of interaction of compression shock with boundary layer. Shockless pressure distributions at supercritical Mach numbers were found to be accounted for by a marked thickening of the boundary layer for some distance ahead of a shock wave.
Report presenting the results of testing in the transonic range of four flutter airfoils attached to a freely falling body. Failures of the airfoils were metered and recorded in order to determine the Mach numbers and altitudes of failure.
Report presenting the results of an experimental investigation of the aerodynamic characteristics of a rotating axial-flow blade grid with pressure-increasing effect. Several techniques of measurement were used, including pressure distribution measurements, pitot tubes, and hot wire wake surveys. Results regarding the lift and pressure drag of the blade sections, aerodynamic characteristics of the blade sections, profile drag, and static pressure are provided.
Investigations were conducted to determine effectiveness of refrigerants in increasing thrust of turbojet engines. Mixtures of water an alcohol were injected for a range of total flows up to 2.2 lb/sec. Kerosene was injected into inlets covering a range of injected flows up to approximately 30% of normal engine fuel flow. Injection of 2.0 lb/sec of water alone produced an increase in thrust of 35.8% of rate engine conditions and kerosene produced a negligible increase in thrust. Carbon dioxide increased thrust 23.5 percent.
Pressure distribution over an extended leading-edge flap on a 42 degree swept-back wing was investigated. Results indicate that the flap normal-force coefficient increased almost linearly with the angle of attack to a maximum value of 3.25. The maximum section normal-force coefficient was located about 30 percent of the flap span outboard of the inboard end and had a value of 3.75. Peak negative pressures built up at the flap leading edge as the angle of attack was increased and caused the chordwise location of the flap center of pressure to be move forward.
Report presenting an investigation in the two-dimensional low-turbulence tunnel and low-turbulence pressure tunnel to determine the highest maximum lift configurations of a slotted flap on an NACA 65(sub 112)A111 (approx.) airfoil section. The scale effects on aerodynamic characteristics were determined for a range of Reynolds numbers. Results regarding flap configurations, lift characteristics, pitching-moment characteristics, drag characteristics, and effects of leading-edge roughness are provided.
Report presenting pressure-distribution measurements made in steady straight and accelerated flight over a horizontal tail surface of a typical pursuit airplane. Results regarding the variation of horizontal-tail loads with Mach number, variation of root bending moments and fuselage torsional moments with Mach number, and variation of section loads with Mach number are provided.
An approximation method for three-dimensional axially symmetrical supersonic flows is developed; it is based on the characteristics theory (represented partly graphically, partly analytically). Thereafter this method is applied to the construction of rotationally symmetrical nozzles. (author).
Report presenting a method for predicting the stability of automatically controlled aircraft by comparing the calculated frequency-response curves for the aircraft and experimentally determined frequency-response curves for the automatic pilot. The method is applicable only to stabilization in roll. It can be used to establish the specifications of performance required for the automatic control device for pilotless aircraft designed as missiles.
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.
Report presenting a method of designing vaneless diffusers using data given for simple conical diffusers. A family of diffusers with three different cone angles and the same throat height was designed and experimentally studied. A second set of diffusers with varying throats with the best cone angle was also investigated.
From Summary: "A method, based on the elastic theory for plate buckling and test results for 24S-T aluminum-alloy Z-stiffened panels, is shown for calculating the compressive strength of Z-stiffened panels that develop local instability. This method can be used to calculate the critical compressive stress above, as well as within, the elastic range. For stresses above three-fourths the compressive yield stress the method can be used for the approximate determination of the average compressive stress at maximum load."
From Summary: "An investigation has been conducted on a V-1650-7 engine to determine the cylinder temperatures and the coolant and oil heat rejections over a range of coolant flows (50 to 200 gal/min) and oil inlet temperatures (160 to 2150 F) for two values of coolant outlet temperature (250 deg and 275 F) at each of four power conditions ranging from approximately 1100 to 2000 brake horsepower. Data were obtained for several values of block-outlet pressure at each of the two coolant outlet temperatures. A mixture of 30 percent by volume of ethylene glycol and 70-percent water was used as the coolant."
From Summary: "An investigation was conducted to determine the coolant-flow distribution, the cylinder temperatures, and the heat rejections of the V-1650-7 engine . The tests were run a t several power levels varying from minimum fuel consumption to war emergency power and at each power level the coolant flows corresponded to the extremes of those likely to be encountered in typical airplane installations, A mixture of 30-percent ethylene glycol and 70-percent water was used as the coolant. The temperature of each cylinder was measured between the exhaust valves, between the intake valves, in the center of the head, on the exhaust-valve guide, at the top of the barrel on the exhaust side, and on each exhaust spark-plug gasket."
From Summary: "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 l0-(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 20deg to 55deg at various rotational speeds and at airspeeds up to 496 miles per hour. The resultant tip speeds obtained simulate actual flight conditions, and the variation of air-stream Mach number with advance ratio is within the range of full-scale constant-speed propeller operation."
From Summary: "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."
A series of 11 fuels ranging in volatility and including various types of hydrocarbons were tested in a single tubular combustion chamber of a turbojet engine under inlet-air conditions simulating engine operation at two speeds at an altitude of 40,000 feet. Temperature-rise data at various fuel-air ratios were obtained for each set of air-flow conditions. Results regarding the effect of combustor inlet-air conditions on temperature rise, four different series of tests, and a review of some general considerations are provided.
Report presenting the performance of a set of axial-flow fan and compressor rotor blades with high design loading in a low-speed test blower. The efficiency curve, efficiency contours, comparison of pitch-section turning angles, and simulated blade roughness are provided.
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