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 Collection: Technical Report Archive and Image Library
Tests of the Northrop XB-62 missile in the Ames 40 by 80 foot wind tunnel
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Tests of the Northrop XSSM-A-3 Missile in the Ames 40- by 80-Foot Wind Tunnel: Wing Modifications
Wind-tunnel tests were conducted to determine the longitudinal stability characteristics of a full-scale Northrop XSSM-A-3 missile. Various wing modifications were investigated in an effort to provide a configuration that would maintain longitudinal stability to lift coefficients necessary for landing the missile during flight tests. The results of the tests led to the choice of a wing with an increased leading-edge radius. A short discussion of the results is presented, but no analysis of the data has been made in order to make the data available as soon as possible.
Tests of Thermal-Electric De-Icing Equipment for Propellers
Flights were made in natural icing conditions at the NACA Ice Research Project, Minneapolis, Minn. to test several designs of thermal-electric propeller de-icing blade shoes and a hub-generator design. It was found that a minimum average unit power of 2.5 watts per square inch of blade-shoe area would protect the propeller blades at the test conditions. The most satisfactory blade shoe of the three designs tested extended to the 20-percent-chord point and to 90 percent of the blade radius. A concentration of heat in the leading-edge region of this shoe was found to reduce the power input necessary for satisfactory de-icing. A satisfactory thermal design of blade shoe and a hub generator of sufficient capacity were developed.
Tests of three tapered airfoils based on the N.A.C.A. 2200, the N.A.C.A.-M6, and the Clark Y sections
Three tapered airfoils based on the N.A.C.A. 2200, the N.A.C.A.-M6, and the Clark Y sections were tested in the variable-density wind tunnel at a Reynolds Number of approximately 3,100,000. The models, which were of aspect ratio 6, had constant core center sections and rounded tips, and tapered in thickness from 18 percent at the roots to 9 percent at the tips. The aerodynamic characteristics are given by the usual dimensionless coefficients plotted for both positive and negative angles of attack and by effective profile-drag coefficients plotted against lift coefficients.
Tests of Two-Blade Propellers in the Langley 8-Foot High-Speed Tunnel to Determine the Effect on Propeller Performance of a Modification of Inboard Pitch Distribution
Tests of two propellers having two blades and differing only in the inboard pitch distribution were made in the Langley 8-foot highspeed tunnel to determine the effect of inboard pitch distribution on propeller performance. propeller was designed for operation in the reduced velocity region ahead of an NACA cowling; the inboard pitch distribution of the modified propeller was increased for operation at or near free-stream velocities, such as would be obtained in a pusher installation. conditions covering climb, cruise, and high-speed operation. Wake surveys were taken behind the propellers in order to determine the distribution of thrust along the blades and to aid in the analysis of the results. Test results showed that the modified propeller was about 2.5 percent less efficient for a typical climb condition at all altitudes, 2 percent more efficient for one cruise condition, and 5 percent more efficient for high-speed operation. speed condition, the modified propeller showed a 6-percent loss in efficiency due to compressibility; whereas the original propeller showed an 11-percent efficiency loss due to compressiblity. The lower compressibility loss for the modified propeller resulted from the fact that the inboard sections of this propeller could operate at increased thrust loading after compressibility losses had occurred at the outboard sections.
Tests of two full-scale propellers with different pitch distributions, at blade angles up to 60 degrees
Two 3-blade 10-foot propellers were operated in front of a liquid-cooled engine nacelle. The propellers differed only in pitch distribution; one had normal distribution (nearly constant pitch for a blade angle of 15 degrees at 0.75 radius), and the other had the pitch of the tip sections decreased with respect to that for the shank sections (blade angle of 35 degrees for nearly constant pitch distribution). Propeller blade angles at 0.75r from 15 degrees to sixty degrees, corresponding to design speeds up to 500 miles per hour, were investigated. Propeller blade angles at 0.75r from 15 degrees to 60 degrees, corresponding to design speeds up to 500 miles per hour, were investigated. The results indicated that the propulsive efficiency at a blade angle of 60 degrees was about 9 percent less than the maximum value of 86 percent, which occurred at blade angle of about 30 degrees. The efficiency at a blade angle of 60 degrees was increased about 7 percent by correcting for the effect of a spinner and, at a blade angle of 30 degrees about 3 percent. The peak efficiencies for the propeller having the washed-out pitch distribution were slightly less than for the normal propeller but the take-off efficiency was generally higher.
Tests of two models representing intermediate inboard and outboard wing sections of the XB-36 airplane
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Tests of Wing Machine-Gun and Cannon Installations in the NACA Full-Scale Wind Tunnel, Special Report
At the request of the Bureau of Aeronautics, an investigation was conducted in the full-scale wind tunnel of wing installations of .50-caliber machine guns and 20-millimeter cannons. The tests were made to determine the effect of various gun installations on the maximum lift and the high-speed drag of the airplane.
Tests on a model of the D.F.W. airplane T-29 of the "Deutsche Flugzeug derke" (German airplane works)
Experiments similar to those carried out with the A.E.G (Allgemeine Elektrizitats-Gesellschaft) were made in the small wind tunnel of the Gottingen laboratory on a model of the D.F.W. airplane T-29. Three series of tests were carried out on the model with a velocity head (or dynamic pressure) of 5 kg/sq m (1.02 lb/sq ft), during which one of the movable surfaces was deflected at various angles, while both the others were retained in their central positions. Of special interest among the results of the tests is the different run of the elevating moments. The curves for the A.E.G. model, rising to the right, denote stability with the elevator locked, while the slight inclination to the left with the D.F.W model denotes a slight instability. For the maximum C(sub L) values, the stability of A.E.G. model continues to increase and the instability of the D.F.W. model is converted into stability. The rolling moments shown when the angular deflection of the ailerons is 0 degrees are due, in both series of tests, to the unequal distribution of the air velocity over the cross section of the wind tunnel, rather than to a lack of symmetry in the model.
Tests on air propellers in yaw
This report contains the results of tests to determine the thrust (pull) and torque characteristics of air propellers in movement relative to the air in a line oblique to the line of the shaft, and specifically when such angle of obliquity is large, as in the case of helicopter flight with the propeller serving for both sustentation and traction.
Tests on airplane fuselages, floats and hulls
This report is a compilation of test data on airplane fuselages, nacelles, airship cars, seaplane floats, and seaplane hulls, prepared by the Bureau of Aeronautics, at the requests of the National Advisory Committee for Aeronautics. The discussion of the data includes the derivation of a scale correction curve to be used in obtaining the full scale drag. Composite curves of drag and L/D for floats and hulls are also given. (author).
Tests on an airfoil with two slots suitable for an aircraft of high performance : lift, drag, rolling and yawing moment measurements
The results that are described in this article form a complete series of tests on an airfoil fitted with front and rear slots, the rear slot being formed between the portion of the wing aft of the rear spar and the forward portion of the flap.
Tests on an airplane model, AEG D I of the Allgemeine Elektricitats Gesellschft, A-G, airplane construction section conducted at the Gottingen Model Testing Laboratory for Aerodynamics
Tests were carried out in the small wind tunnel of the Gottingen establishment on a complete model of the AEG D I airplane. The agreement between the model and the complete airplane applies particularly to the wings, which have ribs cut out of sheet metal and built up in exactly the same manner as in the actual airplane. Various series of tests were carried out with this model in which one or the other of the control surfaces were adjusted to various angles, while the others remained in their neutral positions. During the first three series of tests, the stabilizer was set at a positive angle of 3 degrees, 45' relative to the axis of the engine crankshaft, after which further tests at a 6 degree 30' we made. Finally, the model was tested with the tail group removed. With the elevators set in the prescribed positions, the lift, the drag, and the moments about an axis passing through the center of gravity and perpendicular to the plane of symmetry were measured. All three sets of readings are given as absolute coefficients. Where one of the other control surfaces was deflected from its normal position, the moment produced by that adjustment of the surface was also measured.
Tests on built-up airplane struts having initial tension in outside fibers
The conventional airplane strut fails by buckling. The first signs of failure usually are compression cracks on the concave side of bending. It would therefore appear that if an initial tension could be introduced in the outside fibers, this tension would have to be relieved before the compression load could make itself felt.
Tests on duralumin columns for aircraft construction
The following paper is based on the results of tests, upon duralumin columns, contained in two theses presented to the Department of Civil and Sanitary Engineering of the Massachusetts Institute of Technology.
Tests on models of three British airplanes in the variable density wind tunnel
This report contains the results of tests made in the National Advisory Committee for Aeronautics variable density wind tunnel on three airplane models supplied by the British Aeronautical Research Committee. These models, the BE-2E with R.A.F. 19 wings, the British Fighter with R.A.F. 15 wings, and the Bristol Fighter with R.A.F. 30 wings, were tested over a wide range in Reynolds numbers in order to supply data desired by the Aeronautical Research Committee for scale effect studies. The maximum lifts obtained in these tests are in excellent agreement with the published results of British tests, both model and full scale. No attempt is made to compare drag data, owing to the emission of tail surfaces, radiator, etc., from the model, but is shown that the scale effect observed on the drag coefficients in these tests is due to a large extent to the parts of the models other than the wings. (author).
Tests on riveted joints in sheet duralumin
In making tension tests, the slippage of the joints was noted at three points across each joint. In addition, stress strain curves were obtained for plain tension specimens, and a chemical analysis was made of the sheet.
Tests on stiffened circular cylinders
Compressive tests were made of two series of stiffened circular cylindrical shells under axial load. All the shells were 16 inches in diameter by 24 inches in length and were made of aluminum-alloy sheet curved to the proper radius and welded with one longitudinal weld. The ratios of diameter to thickness of shell wall in the two series of specimens were 258 and 572. Strains were measured with Huggenberger tensometers at a number of gage lines on the stiffeners and shell. The results of these tests indicate that a spacing of circumferential stiffeners equal to 0.67 times the radius is too great to strengthen the shell wall appreciably. The results are not inclusive enough to show the optimum in stiffeners. Plain cylinders without stiffeners developed ultimate strengths approximately half as great as the buckling strengths computed by the equation resulting from the classical theory and slightly greater than those computed by Donnell's large deflection theory.
Tests on thin-walled celluloid cylinders to determine the interaction curves under combined bending, torsion, and compression or tension loads
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Tests on thirteen navy type model propellers
The tests on these model propellers were undertaken for the purpose of determining the performance coefficients and characteristics for certain selected series of propellers of form and type as commonly used in recent navy designs. The first series includes seven propellers of pitch ratio varying by 0.10 to 1.10, the area, form of blade, thickness, etc., representing an arbitrary standard propeller which had shown good results. The second series covers changes in thickness of blade section, other things equal, and the third series, changes in blade area, other things equal. These models are all of 36-inch diameter. Propellers A to G form the series on pitch ratio, C, N. I. J the series on thickness of section, and K, M, C, L the series on area. (author).
Tests on thrust augmenters for jet propulsion
This series of tests was undertaken to determine how much the reaction thrust of a jet could be increased by the use of thrust augmenters and thus to give some indication as to the feasibility of jet propulsion for airplanes. The tests were made during the first part of 1927 at the Langley Memorial Aeronautical Laboratory. A compressed air jet was used in connection with a series of annular guides surrounding the jet to act as thrust augmenters. The results show that, although it is possible to increase the thrust of a jet, the increase is not large enough to affect greatly the status of the problem of the application of jet propulsion to airplanes.
Tests to determine effects of slipstream rotation on the lateral stability characteristics of a single-engine low-wing aircraft model
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Tests to Determine the Adhesive Power of Passenger-Car Tires
The concept of the adhesive power of a tire with respect to the road involves several properties which result from the purpose of the tire; namely, connecting link between vehicle and road: (1) The tire must transfer the tractive and braking forces acting in the direction of travel (tractive and braking adhesion); (2) The tire is to prevent lateral deviations of the vehicle from the desired direction of travel (track adhesion). Moreover, the rubber tire provides part of the springing of the vehicle. Above all, it has to level out the minor road irregularities; thus it smoothes, as it were, the road and simultaneously reduces the noise of driving. The springing properties of the tire affect the adhesive power. The tests described below comprise a determination of the braking and track adhesion of individual tires. The adhesion of driven wheels has not been investigated so far.
Tests to determine the effect of heat on the pressure drop through radiator tubes
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Tests with hydrogen fuel in a simulated afterburner
Hydrogen fuel combustion analysis in simulated afterburner.
Tests with hydrogen fuel in a simulated afterburner
An investigation was conducted in a 16-inch-diameter simulated afterburner using gaseous hydrogen fuel. No flameholder was used with a multipoint fuel injector. The burner length was varied from 9.5 to 38 inches. The afterburner-inlet conditions were: temperature of 1200 degrees or 1500 degrees F, pressure of 14 to 44 inches mercury absolute, and velocity of 300 to The measured combination efficiency ranged from 85 to 98 percent over an equivalence-ratio range of 0.2 to 1.0. The cold-flow pressure-drop coefficient was 1.0 for the system. Spontaneous ignition was always possible at temperatures above 1200 degrees F but was not possible at temperatures above 1200 degrees F but was not possible below 1100 degrees F for all pressures and velocities tested.
Theoretical additional span loading characteristics of wings with arbitrary sweep, aspect ratio, and taper ratio
The Weissinger method for determining additional span loading has been used to find the lift-curve slope, spanwise center of pressure, aerodynamic center location, and span loading coefficients of untwisted and uncambered wings having a wide range of plan forms characterized by various combinations of sweep, aspect ratio, and taper ratio. The results are presented as variations of the aerodynamic characteristics with sweep angle for various values of aspect ratio and taper ratio. Methods are also included for determining induced drag and the approximate effects of compressibility. Despite the limitations of a lifting line method such as Weissinger's, the good agreement found between experimentally and theoretically determined characteristics warrants confidence in the method. In particular, it is believed that trends observed in results of the Weissinger method should be reliable. One of the most significant results showed that for each angle of sweep there is a taper ratio for which aspect ratio has little effect on the span loading and for which the loading is practically elliptical. This elliptic loading is approached at a taper ratio of 1.39 for 30 degree of sweepforward, 0.45 for zero degree of sweepback. (author).
Theoretical aerodynamic characteristics of a family of slender wing-tail-body combinations
The aerodynamic characteristics of an airplane configuration composed of a swept-back, nearly constant chord wing and a triangular tail mounted on a cylindrical body are presented. The analysis is based on the assumption that the free-stream Mach number is near unity or that the configuration is slender. The calculations for the tail are made on the assumption that the vortex system trailing back from the wing is either a sheet lying entirely in the plane of the flat tail surface or has completely "rolled up" into two point vortices that lie either in, above, or below the plane of the tail surface.
Theoretical aerodynamic characteristics of bodies in a free-molecule-flow field
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Theoretical aerodynamic coefficients of two-dimensional supersonic biplanes
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Theoretical analyses to determine unbalanced trailing-edge controls having minimum hinge moments due to deflection at supersonic speeds
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Theoretical analyses to determine unbalanced trailing-edge controls having minimum hinge moments due to deflection at supersonic speeds
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Theoretical Analysis and Bench Tests of a Control-Surface Booster Employing a Variable Displacement Hydraulic Pump
The NACA is conducting a general investigation of servo-mechanisms for use in powering aircraft control surfaces. This paper presents a theoretical analysis and the results of bench tests of a control-booster system which employs a variable displacement hydraulic pump. The booster is intended for use in a flight investigation to determine the effects of various booster parameters on the handling qualities of airplanes. Such a flight investigation would aid in formulating specific requirements concerning the design of control boosters in general. Results of the theoretical analysis and the bench tests indicate that the subject booster is representative of types which show promise of satisfactory performance. The bench tests showed that the following desirable features were inherent in this booster system: (1) No lost motion or play in any part of the system; (2) no detectable lag between motion of the contra1 stick and control surface; and (3) Good agreement between control displacements and stick-force variations with no hysteresis in the stick-force characteristics. The final design configuration of this booster system showed no tendency to oscillate, overshoot, or have other undesirable transient characteristics common to boosters.
A theoretical analysis of a simple aerodynamic device to improve the longitudinal damping of a cruciform missile configuration at supersonic speeds
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Theoretical analysis of an airplane acceleration restrictor controlled by normal acceleration, pitching acceleration, and pitching velocity
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A theoretical analysis of elastic vibrations of fixed-ended and hinged helicopter blades in hovering and vertical flight
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A theoretical analysis of heat transfer in regions of separated flow
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Theoretical analysis of hydrodynamic impact of a prismatic float having freedom in trim
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Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows
A method for the solution of the incompressible nonviscous flow through a centrifugal impeller, including the inlet region, is presented. Several numerical solutions are obtained for four weight flows through an impeller at one operating speed. These solutions are refined in the leading-edge region. The results are presented in a series of figures showing streamlines and relative velocity contours. A comparison is made with the results obtained by using a rapid approximate method of analysis.
Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows I : solution by a matrix method and comparison with an approximate method
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Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows II : solution in leading-edge region by relaxation methods
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Theoretical analysis of linked leading-edge and trailing-edge flap-type controls at supersonic speeds
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Theoretical analysis of one-stage windmills for reducing flow distortion
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Theoretical analysis of oscillations in hovering of helicopter blades with inclined and offset flapping and lagging hinge axes
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Theoretical analysis of oscillations of a towed cable
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Theoretical analysis of some simple types of acceleration restrictors
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Theoretical Analysis of Stationary Potential Flows and Boundary Layers at High Speed
The present report consists of two parts. The first part deals with the two-dimensional stationary flow in the presence of local supersonic zones. A numerical method of integration of the equation of gas dynamics is developed. Proceeding from solutions at great distance from the body the flow pattern is calculated step by step. Accordingly the related body form is obtained at the end of the calculation. The second part treats the relationship between the displacement thickness of laminar and turbulent boundary layers and the pressure distribution at high speeds. The stability of the boundary layer is investigated, resulting in basic differences in the behavior of subsonic and supersonic flows. Lastly, the decisive importance of the boundary layer for the pressure distribution, particularly for thin profiles, is demonstrated.
A theoretical analysis of the distortion of fuel-spray-particle paths in a helicopter ram-jet engine due to centrifugal effects
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A theoretical analysis of the effect of aileron inertia and hinge moment on the maximum rolling acceleration of airplanes in abrupt aileron rolls
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A theoretical analysis of the effect of engine angular momentum on longitudinal and directional stability in steady rolling maneuvers
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