From Introduction: "The present paper presents the scale effect on the longitudinal aerodynamic characteristics, the aerodynamic characteristics in yaw, and the tuft studies for 0^o and 3.7^o yaw. The results of the effect of leading-edge and trailing-edge flaps on the aerodynamic characteristics of the wing will be presented in later reports."
From Introduction: "Evaluation of tail-pipe burning in this engine with a larger tail-pipe combustion chamber is discussed in reference 1. Results of investigations on tail-pipe burning in this engine at static sea-level conditions are presented in reference 2. An investigation of thrust augmentation by means of injecting water at the inlet of an axial-flow compressor engine is discussed in reference 3."
Report presenting a combustion-efficiency investigation of 10 special straight-run distillate fuels in an individual tubular-type combustor unit of a 14-unit assembly at two simulated engine operating conditions. These distillates were obtained from various crude oils and consisted of hydrocarbon mixtures with distillations temperature varying from 93 to 690 degrees Fahrenheit. Comparison of temperature measurements obtained from two locations in the exhaust duct showed that under certain operating conditions the flame extended beyond the turbine position.
Report presenting a flight investigation of a scout-bomber airplane with 4.7 degrees and 10 degree geometric dihedral in the wing outer panels in order to obtain flight test results pertaining to the upper limit of the wing-dihedral angle for satisfactory handling qualities. No unusual or undesirable lateral control characteristics were observed with the combinations of stability parameters obtained with the two dihedral angles used.
The evaporation velocity of liquid droplets under various conditions is theoretically calculated and a number of factors are investigated which are neglected in carrying out the fundamental equation of Maxwell. It is shown that the effect of these factors at the small drop sizes and the small weight concentrations ordinarily occurring in fog can be calculated by simple corrections. The evaporation process can be regarded as quasi-stationary in most cases. The question at hand, and also the equivalent question of the velocity of growth of droplets in a supersaturated atmosphere, is highly significant in meteorology and for certain industrial purposes. Since the literature concerning this is very insufficient and many important aspects either are not considered at all or are reported incorrectly, it seems that a short discussion is not superfluous. A special consideration will be given to the various assumptions and neglections that are necessary in deriving the fundamental equation of Maxwell. The experimental work available, which is very insufficient and in part poorly dependable, can be used as an accurate check on the theory only in very few cases.
In the case of cones in axially symmetric flow of supersonic velocity, adiabatic compression takes place between shock wave and surface of the cone. Interpolation curves betwen shock polars and the surface are therefore necessary for the complete understanding of this type of flow. They are given in the present report by graphical-numerical integration of the differential equation for all cone angles and airspeeds.
Report presenting design charts for 24S-T and 75S-T aluminum-alloy flat compression panels with longitudinal extruded Y-section stiffeners. Comparisons are also made among panels designed from these charts and 24S-T aluminum-alloy panels with formed Z-section stiffeners designed from available design charts.
Stress-rupture data for four heat-resisting alloys are analyzed according to equations of the theory of rate processes. A method for determining the four parameters of structure and composition is demonstrated and the four parameters are determined for each of the alloys: forged S816, cast S816, cast S590, and cast Vitallium. It is concluded that parameters can be determined for an alloy provided sufficient reliable experimental data are available.
The results of an experimental investigation made for the purpose of developing suitable jet-engine nacelle designs for a high-speed medium bomber are presented. Two types of nacelles were investigated, the first enclosing two 4000-pounds-thrust jet engines and a 65-inch-diameter landing wheel and the second enclosing a single 4000-pounds-thrust jet engine. Both types of nacelles were tested in positions underslung beneath the wing and centrally located on the wing. This report summarizes the investigation which was performed at low speed for the purpose of developing entrance and body shapes of suitable form. Included are results from the high-speed portion of the investigation on the characteristics of an underslung nacelle.
Report presenting the results of high-speed wind-tunnel research on the effects of modifications to the horizontal tail profile on the static longitudinal stability and control of a pursuit airplane at high speeds. Two symmetrical stabilizers, two flat-sided elevators, and three elevators with bulged profiles were investigated. The pitching-moment and elevator hinge-moment characteristics with various tails are shown.
From Summary: "An investigation was made to determine the effects of changes in the amount and distribution of forebody and afterbody dead rise on the hydrodynamic resistance and spray characteristics of a 1/11-size model of the Bureau of Aeronautics design No. 22ADR class VPB airplane. The variations in dead rise within the range investigated had no significant effects on resistance or trim, free to trim, or on resistance or trimming moment, fixed in trim. The coordinates of the peaks of the bow-spray blisters, with reference to the model, were measured at low speeds, and it was found that the model with the low dead rise at the bow had the lowest blisters."
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.
Report presents the results of an investigation conducted to determine some of the effects of airfoil section and washout on the experimental and calculated characteristics of 10-percent-thick wings. Three wings of aspect ratio 9 and ratio of root chord to tip chord 2.5 were tested. One wing had NACA 64-210 sections and 2 degree washout, the second had NACA 65-210 sections and 2 degree washout, and the third had NACA 65-210 sections and 0 degree washout. It was found that the experimental characteristics of the wings could be satisfactorily predicted from calculations based upon two-dimensional data when the airfoil contours of the wings conformed to the true airfoil sections with the same high degree of accuracy as the two-dimensional models.
From Summary: "The knock-limited power at various blend compositions for several aromatics and cycloparaffins individually blended with paraffinic base stocks, as determined in an air-cooled aircraft-engine cylinder at fuel-air ratios of 0.07 and 0.10 is presented. An analysis of the data leads to the conclusion that the extended reciprocal blending relation suggested in a previous NACA report is not generally applicable to such nonparaffinic components, but might possibly be useful as an approximation over a limited range of composition for aromatic blends."
Note presenting data from flight measurements of the meteorological factors related to the intensity of icing conditions. The physical factors that establish the distribution of liquid water in clouds of various types are discussed and the results of the analysis are used to formulate certain rules for the forecasting of icing intensity. Discussion of the results includes the problem of forecasting icing conditions and the problem of defining the physical characteristics of icing conditions for the purpose of the design of ice-protection equipment.
Note presenting results of a flight investigation conducted on a fighter-type airplane to determine the factors which affect the loads and load distributions on the vertical tail surfaces in maneuvers. An analysis is made of the data obtained in steady flight, rudder kicks, and fishtail maneuvers.
A wind-tunnel investigation of the Boeing XB-47 full-scale empennage was conducted to provide, prior to flight tests, data required on the effectiveness of the elevator and rudder. The XB-47 airplane is a jet-propelled medium bomber having wing and tail surfaces swept back 35 degrees. The investigation included tests of the effectiveness of the elevator with normal straight sides, with a buldged trailing edge, and with a modified hinge-line gap and tests of the effectiveness of the rudder with a normal straight-sided tab and with a bulged tab.
Report presenting an aerodynamic-control-effectiveness investigation using free-flight rocket-propelled RM-5 test vehicles at high subsonic, transonic, and supersonic speeds. Results regarding aileron control characteristics and drag measurements are provided.
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.
"Measurements have been made of the mean-total-head and mean-temperature fields in a round turbulent jet with various initial temperatures. The results show that the jet spreads more rapidly as its density becomes lower than that of the receiving medium, even when the difference is not sufficiently great to cause dynamic-pressure function. Rough analytical considerations have given the same relative spread. The effective "turbulent Prandtl number" for a section of the fully developed jet was found to be equal to the true (laminar) Prandtl number within the accuracy measurement" (p. 859).
"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 a scout-bomber 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" (p. 1).
Report presenting a high-speed wind-tunnel investigation of a pitot-static tube to provide information on the effects of compressibility upon the pressure indications of a representative airspeed head at high subsonic speeds. The calibration factor for the instrument has been evaluated for several small angles of pitch and yaw throughout a Mach number range from 0.30 to approximately 0.925. The results indicate that the calibration factor for each combination of pitch and yaw angles tested is almost constant with Mach number up to a Mach number of approximately 0.8.
Report presenting an investigation to determine the effects of compressibility on the aerodynamic characteristics of a 5-inch-chord NACA 16-009 airfoil with a 32.9-percent-chord flap with a nose overhang and an unsealed gap. Airfoil lift and pitching moment and flap hinge moments were obtained for a range of angles of attack, flap angle ranges, and Mach numbers.
Report presenting the results of an investigation of a specific tail configuration in the 15-foot free-spinning tunnel in order to supplement the existing published data on hinge moments of elevators and rudders in spins. Hinge-moment measurements are presented for a balanced elevator equipped with trim tabs and for a balanced rudder. Results regarding elevator hinge moments, rudder hinge moments, and application of hinge-moment data are provided.
Report presenting an investigation to obtain a record of any chemical changes directly associated with knock in a spark-ignition engine on a greatly magnified time scale by means of coordinated high-speed motion pictures and spectrograms.
Report presents the results of an investigation conducted in the Langley 19-foot pressure tunnel to determine the maximum lift and stalling characteristics of two thin wings equipped with several types of flaps. Split, single slotted, and double slotted flaps were tested on one wing which had NACA 65-210 airfoil sections and split and double slotted flaps were tested on the other, which had NACA 64-210 airfoil sections. Both wings were zero sweep, an aspect ratio of 9, and a taper ratio of 0.4.
"Tests of a partial-span model of a large bomber-type airplane were conducted to determine the aerodynamic characteristics of the wing equipped with full-span flaps and a retractable spoiler end aileron lateral control system. The arrangement consisted of (1) a double slotted flap extending over aproximate1y 86 percent of the wing semispan, (2) a 20-percent constant-percentage-chord aileron extending from the outboard end of the flap to the wing tip, and (3) a retractable spoiler, located at the 65-percent wing-chord station and extending from approximately 63 percent of the wing semispan to the wing tip. In addition, tests were made of a wing vent (of 1 and 2 percent of the wing chord located directly behind the spoiler), perforations in the spoiler, a blot or cut-out along the lower edge of the spoiler and spoilers of various spans" (p. 1).
An investigation is in progress at the Langley Laboratory of the NACA to explore the possibilities of axial-flow compressors operating with supersonic velocities relative to the blade rows. The first phase of this investigation, a study of supersonic diffusers, has been reported. The second phase, an analysis of supersonic compressors, has also been reported. Preliminary calculations have shown that very high pressure ratios across a stage, together with somewhat increased mass flows, are possible with compressors which decelerate air through the speed of sound in their rotor blading. These performance characteristics are desirable in compressors for aircraft jet propulsion units, gas turbines, or superchargers. The third phase, presented here, is a preliminary experimental investigation of a supersonic compressor designed to produce a high pressure ratio in a single stage.
The sea-level performance of I-16 turbojet engine at zero ram was investigated to determine the effects of an intake duct, shroud, and tail pipe intended for installation in an XFR-1 airplane. Engine speeds ranged from 8000 to 16,500 rpm for several variations of the intake duct and tail pipes.
Report presenting total-pressure recoveries obtained with four-cone inlet combinations at Mach number 1.85. The configurations tested included a cone designed to produce three oblique shocks ahead of the diffuser inlet combined with two other inlets, a cone generated by a parabolic arc in combination with two other inlets, a cone-inlet combination designed to produce an isentropic entrance flow at 0 degrees angle of attack, and a 30 degree single-shock cone combined with a perforated inlet section. Each of the configurations yielded total-pressure recoveries higher than what was reported in previous testing.
Tests have been conducted in the Langley high-speed 7- by 10-foot tunnel over a Mach number range from 0.40 to 0.91 to determine the stability and control characteristics of an 0.08-scale model of the Chance Vought XF7U-1 airplane. The aileron characteristics of the complete model are presented in the present report with a very limited analysis of the results.
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.
Effect of inlet-air pressure and temperature on the performance of the X24-2 10-Stage Axial-Flow Compressor from the X24C-2 turbojet engine was evaluated. Speeds of 80, 89, and 100 percent of equivalent design speed with inlet-air pressures of 6 and 12 inches of mercury absolute and inlet-air temperaures of approximately 538 degrees, 459 degrees,and 419 degrees R ( 79 degrees, 0 degrees, and minus 40 degrees F). Results were compared with prior investigations.
Note presenting a discussion of the integral relations for flow of the boundary-layer type. It is shown that the characteristic laws of spread of jets, wakes, and so forth, can be obtained directly for the laminar case and, with the help of dimensional reasoning, for the turbulent case as well. Measurements of the mean velocity, the intensity and scale of the turbulent fluctuations, and of the turbulent shear in a two-dimensional mixing zone are presented.
Memorandum presenting investigation on reductions of friction on wings, especially by means of boundary-layer suction. The report is broken up into several sections, including: causes of transition, laminar profiles with the transition taking place, laminar boundary-layer suction, investigation of the laminar pressure increase, investigation of the slot flow for laminar boundary-layer suction with single slots, tests about keeping a boundary layer for high Reynolds laminar with the aid of boundary-layer suction, and an investigation of a slightly-cambered laminar suction profile of 10.5-percent thickness.
From Summary: "An investigation of the DM-1 Glider, which had approximately triangular plan form, an aspect ratio of 1.8 and a 60 degree sweptback leading edge, has been conducted in the Langley full-scale tunnel. The investigation consisted of the determination of the separate effects of the following modifications made to the glider on its maximum lift and stability characteristics: (a) installation of sharp leading edges over the inboard semispan of the wing, (b) removal of the vertical fin, (c) sealing of the elevon control-balance slots, (d) installation of redesigned thin vertical surfaces, (e) installation of faired sharp leading edges, and (f) installation of canopy."
The lift coefficient of a wing of small span at first shows a linear increase for the increasing angle of attack, but to a lesser degree then was to be expected according to the theory of the lifting line; thereafter the lift coefficient increases more rapidly than linearity, as contrasted with the the theory of the lifting line. The induced drag coefficient for a given lift coefficient, on the other hand, is obviously much smaller than it would be according to the theory. A mall change in the theory of the lifting line will cover these deviations.
Report presenting measurements using the NACA wing-flow method of the lift, pitching-moment, and hinge-moment characteristics of a 35 degree sweptback NACA 65-009 airfoil of aspect ratio 3.04 with a full-span 1/4-chord unsealed plain flap. The tests were carried out at a range of Mach numbers, Reynolds numbers, angles of attack, and flap deflections. Variations of lift and pitching moment with angle of attack or flap deflection were approximately rectilinear at all Mach numbers for moderate angles of attack and flap deflections.
Report presenting a method for determining, by step-by-step integration, the trajectories of water drops around any body in two-dimensional flow for which the streamline velocity components are known or can be computed. The equations are presented in general form and then water-drop trajectories are calculated about a 12-percent-thick symmetric Joukowski profile.
Note presenting a method for determining, by step-by-step integration, the trajectories of water drops around any body in two-dimensional flow for which the streamline velocity components are known or can be computed. The equations are presented in general form and then, to illustrate the procedure, water-drop trajectories are calculated about a 12-percent-thick symmetrical Joukowski profile chosen to simulate an NACA 0012 section. The method provides a means for the relatively rapid calculation of the trajectory of a single drop without the utilization of a differential analyzer.
Report presenting a method for the determination by use of charts of mass-flow coefficient and associated flow parameters from pressure surveys in internal-flow systems. The charts presented cover a wide range of specific parameters through the complete range of subsonic Mach numbers. The equations have also been evaluated for flows that involve the adding of mechanical or thermal energy, such as flows behind radiators or propellers.
Report presenting the optical principle of the NACA ultrahigh-speed camera. Simplified sketches are included illustrating the optical principle and main design features of the camera, but without minor design details. A detailed description of the camera is provided as well as some sample photographs taken with the camera.
Note presenting a discussion of the performance characteristics of the compound engine, the turbine-propeller engine, the turbojet engine, the turbo ramjet engine, the ramjet engine, and the rocket engine. Insight is provided into the proper position of each of these engine types in the speed-range spectrum of aircraft operation. It is shown that the compound engine, which has the greatest weight per unit thrust and also the lowest specific fuel consumption, gives the longest range.
Report presenting an investigation to determine whether a safety fuel with a flash point of 122 degrees Fahrenheit could be successfully used in a high-power radial aircraft engine without individual cylinder fuel-injection equipment. The safety fuel was injected into the combustion-air stream in three different ways. Results regarding the standard powers, effect of average fuel-air ratio on mixture distribution, effect of combustion-air temperature on mixture distribution, general engine performance, and idling and starting are provided.
From Summary: "This report presents a rough correlation of the dimensions of water rudders of various actual seaplanes and flying boats as related to their behavior. The correlation should be useful for determining the size of a water rudder which will give adequate control for maneuvering at low speeds."
Investigation at transonic speeds in the high-speed 7- by 10-foot tunnel to determine the rolling-effectiveness characteristics of a spoiler on a double-wedge-type semispan wing with a sweepback angle of 42 degrees. Results regarding the variation of rolling-moment coefficient, spoiler effectiveness, and rolling effectiveness are provided.
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.
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