Report discussing the 6-foot-4-inch wind tunnel and its auxiliary equipment has proven itself capable of continuous and reliable output of data. The real value of the tunnel will increase as experience is gained in checking the observed tunnel performance against full-scale performance. Such has been the case of the 8- by 8-foot tunnel, and for that reason the comparison in the calibration tests have been presented.
This note contains a description of an improved apparatus and procedure used by the NACA for determining the moments of inertia of airplanes. The method used, based on the pendulum theory, is similar to that previously used, but a recent investigation of its accuracy has resulted in the improvements described herein. The error, when using the new apparatus and procedure, has been found to be of the order of 1 per cent.
The various possible means of preventing ice adhesion on airplane surfaces are critically reviewed. Results are presented of tests of the adhesives forces between ice and various solid and liquid forces. It is concluded that the de-icing of airplane wings by heat from engine exhaust shows sufficient promise to warrant full-scale tests. For propellers, at least, and possibly for certain small areas such as windshields, radio masts, etc. the use of de-icing or adhesion-preventing liquids will provide the best means of protection.
Both laboratory and weather-exposure corrosion tests showed conclusively that the protection afforded by aluminum pigmented spar varnish coatings applied to previously anodized aluminum surfaces was greatly superior to that afforded by the same coatings applied to surfaces which had simply been cleaned free from grease and not anodized.
An aerodynamic analysis of the gyroplane rotating-wing system is presented herein. This system consists of a freely rotating rotor in which opposite blades are rigidly connected and allowed to rotate or feather freely about their span axis. Equations have been derived for the lift, the lift-drag ratio, the angle of attack, the feathering angles, and the rolling and pitching moments of a gyroplane rotor in terms of its basic parameters. Curves of lift-drag ratio against lift coefficient have been calculated for a typical case, showing the effect of varying the pitch angle, the solidarity, and the average blade-section drag coefficient. The analysis expresses satisfactorily the qualitative relations between the rotor characteristics and the rotor parameters. As disclosed by this investigation, the aerodynamic principles of the gyroplane are sound, and further research on this wing system is justified.
Tests were made on a model wing with three different sized split trailing-edged flaps, in the NACA 7 by 10 foot wind tunnel. The flaps were formed of the lower rear portion of the wing and were rotated downward about axes at their front edges. The lift, drag, and center of pressure were measured with the axis in its original position and also with it moved back in even steps to the trailing edge of the main wing, giving in effect an increase in area. The split flaps when deflected about their original axis locations gave slightly higher maximum lift coefficients than conventional trailing-edge flaps, and the lift coefficients were increased still further by moving the axes toward the rear. The highest value of C(sub L max), which was obtained with the largest flap hinged at 90 per cent of the chord from the leading edge, was 2.52 as compared with 1.27 for the basic wing.
The effect on airfoil characteristics of surface roughness of varying degrees and types at different locations on an airfoil was investigated at high values of the Reynolds number in a variable density wind tunnel. Tests were made on a number of National Advisory Committee for Aeronautics (NACA) 0012 airfoil models on which the nature of the surface was varied from a rough to a very smooth finish. The effect on the airfoil characteristics of varying the location of a rough area in the region of the leading edge was also investigated. Airfoils with surfaces simulating lap joints were also tested. Measurable adverse effects were found to be caused by small irregularities in airfoil surfaces which might ordinarily be overlooked. The flow is sensitive to small irregularities of approximately 0.0002c in depth near the leading edge. The tests made on the surfaces simulating lap joints indicated that such surfaces cause small adverse effects. Additional data from earlier tests of another symmetrical airfoil are also included to indicate the variation of the maximum lift coefficient with the Reynolds number for an airfoil with a polished surface and with a very rough one.
A number of airfoils, including 14 commonly used airfoils and 10 NACA airfoils, were tested through the negative angle-of-attack range in the NACA variable-density wind tunnel at a Reynolds Number of approximately 3,000,000. The tests were made to supply data to serve as a basis for the structural design of airplanes in the inverted flight condition. In order to make the results immediately available for this purpose they are presented herein in preliminary form, together with results of previous tests of the airfoils at positive angles of attack. An analysis of the results made to find the variation of the ratio of the maximum negative lift coefficient to the maximum positive lift coefficient led to the following conclusions: 1) For airfoils of a given thickness, the ratio -C(sub L max) / +C(sub L max) tends to decrease as the mean camber is increased. 2) For airfoils of a given mean camber, the ratio -C(sub L max) / +C(sub L max) tends to increase as the thickness increases.
The static lift and drag forces on three hemispherical and two conical cups were measured over a range of angles of attack from 0 degrees to 180 degrees and a range of Reynolds Numbers from very small up to 400,000. The problems of supporting the cup for measurement and the effect of turbulence were also studied. The results were compared with those of other investigators.
This paper presents the results of wind-tunnel tests of several airfoils of low aspect ratio. The airfoils included three circular Clark Y airfoils with different amounts of dihedral, two Clark Y airfoils with slots in their portions, and three flat-plate airfoils. Lift, drag, and pitching-moment characteristics of the slotted airfoils with slots open and closed; pitching moment characteristics of one of the slotted airfoils with slots open and closed; and lift characteristics of the flat-plate airfoils are included. The results reveal a definite improvement of lift, drag, and pitching-moment characteristics with increase in dihedral of the circular Clark Y wing. Lift characteristics near the stall were found to depend markedly on the shape of the extreme tip but were not greatly affected by slots through the after portion of the airfoils. Changes in plan form of the flat-plate airfoils gave erroneous indications of the effect to be expected from changes in plan form of an airfoil of Clark Y section. The minimum drag characteristics of the circular Clark Y airfoils were found to be substantially the same as for a Clark Y airfoil of conventional aspect ratio.
This paper presents the results of tests of six commonly used airfoils: the CYH, the N-22, the C-72, the Boeing 106, and the Gottingen 398. The lifts, drags, and pitching moments of the airfoils were measured through a large range of positive and negative angles of attack. The tests were made in the variable density wind tunnel of the National Advisory Committee for Aeronautics at a large Boeing 106, and the Gottingen 398 airfoils, the negative maximum lift coefficients were found to be approximately half the positive; but for the M-6 and the CYH, which have less effective values were, respectively, 0.8 and 0.6 of the positive values.
This report contains the lift, drag, and moment characteristics of tapered Clark Y, Gottingen 393, and USA 45 airfoils as obtained from tests made in the Variable Density Wind Tunnel of the NACA. The results are given at both low and high Reynolds Numbers to show scale effect and to provide data for use in airplane design.
The drag of five models of side floats was measured in the N.A.C.A. 7- by 10-foot wind tunnel. The most promising method of reducing the drag of floats indicated by these tests is lowering the angle at which the floats are rigged. The addition of a step to a float does not always increase the drag in the flying range, floats with steps sometimes having lower drag than similar floats without steps. Making the bow chine no higher than necessary might result in a reduction in air drag because of the lower angle of pitch of the chines. Since side floats are used formally to obtain lateral stability when the seaplane is operating on the water at slow speeds or at rest, greater consideration can be given to factors affecting aerodynamic drag than is possible for other types of floats and hulls.
Measurements of aerodynamic drag were made in the 20-foot wind tunnel on a representative group of 11 flying-boat hull models. Four of the models were modified to investigate the effect of variations in over-all height, contours of deck, depth of step, angle of afterbody keel, and the addition of spray strips and windshields. The results of these tests, which cover a pitch-angle range from -5 to 10 degrees, are presented in a form suitable for use in performance calculations and for design purposes.
Tests were conducted in the N.A.C.A. full scale wind tunnel at the request of the Army Air Corps to determine the effect of retractable landing gear openings in the bottom surface of a wing upon the characteristics of a Lockheed Altair airplane. The tests were extended to include the determination of the lift and drag characteristics throughout the angle-of-attack range with the landing gear both retracted and extended. Covering the wheel openings in the wing with sheet metal when the wheels were extended reduced the drag only 2 percent at a lift coefficient of 1.0, which was assumed for the take-off condition. Therefore, the wheel openings in the bottom side of the wing have a negligible effect upon the take-off of the airplane. Retracting the landing gear reduced the minimum drag of the complete airplane 50 percent.
The investigation described in this report was made to determine the change in aerodynamic forces and moments produced by split flaps in a steady spin. The test were made with the spinning balance in the NACA 5-foot vertical wind tunnel. A low-wing monoplane model was tested with and without the split flaps in 12 spinning attitudes chosen to cover the probable spinning range. The changes in coefficients produced by adding the split flaps are given for longitudinal force, normal force, and rolling and yawing moments about body axes. The results obtained indicate that the use of split flaps on an airplane is unlikely, in any case, to have much beneficial effect on a spin, and it might make the spin dangerous. The change in the spin will depend upon the aerodynamic and inertia characteristics of the particular airplane. A dangerous condition is most likely to be attained with airplanes which are statically stable in yaw in the spinning attitude and which have large weights distributed along wings.
The aerodynamic forces and moments on a 1/12-scale model of the F4B-2 airplane were measured with the spinning balance in nine spinning attitudes with three sets of tail surfaces, namely, F4B-2 surfaces; F4B-4 fin and rudder with rectangular stabilizer; and with all tail surfaces removed. In one of these attitudes measurements were made to determine the effect upon the forces and moments of independent and of simultaneous displacement of the rudder and elevator for two of the sets of tail surfaces. Additional measurements were made for a comparison of model and full-scale data for six attitudes that were determined from flight tests with various control settings. The characteristics were found to vary in the usual manner with angle of attack and sideslip. The F4B-2 surfaces were quite ineffective as a source of yawing moments. The F4B-4 fin and F4B-2 stabilizer gave a greater damping yawing moment when controls were against the spin than did the F4B-2 surfaces but otherwise there was little difference. Substitution of a rectangular stabilizer for the F4B-2 stabilizer made no appreciable difference in the coefficient. Further comparisons with other airplane types are necessary before final conclusions can be drawn as to the relations between model and full-scale spin measurements.
The investigation described in this report was made to determine the effectiveness of floating wing-tip ailerons as an airplane control in the spin. In these tests the ailerons, not being balanced, were set parallel to the axis of rotation, which is probably very nearly the attitude that balanced floating ailerons would assume in a spin. The tests were made with the spinning balance in the N.A.C.A. 5-foot vertical tunnel. The model was tested with and without the ailerons in 12 spinning attitudes chosen to cover the probable spinning range. Rolling- and yawing-moment coefficients are given as measured for the model with and without the ailerons, and computed values are given for the ailerons alone. The addition of floating wing-tip ailerons to the model doubled the rolling-moment coefficient and increased the yawing-moment coefficient by 0.05 and more. Both moments were in a sense to oppose the spin.
Wind tunnel tests were made of a model wing having an aspect ratio of 3 and a tapered plan form with a straight trailing edge. The model had the Clark Y airfoil section throughout it's entire span and had no washout, depending on a trailing-edge flap for longitudinal balance and control. The flap had a constant chord and was divided into four equal portions along the span. The tests were made with the entire flap deflected to obtain longitudinal control and balance, and also with the inner portions deflected alone, and with the outer portions deflected alone. It was found that the simple wing with no washout or change of basic section along the span has aerodynamic characteristics well suited for use on tailless airplanes. A higher lift coefficient was obtained with the full-span flap deflected as a unit to give longitudinal balance than was obtained with either the inner or the outer portions of the flap deflected.
The 'Farnboro' electric indicator was tested as received from the manufacturers, and modifications made to the instrument to improve its operation. The original design of disk valve was altered so as to reduce the mass, travel, and seat area. Changes were made to the recording mechanism, which included a new method of locating the top center position on the record. The effect of friction on the motion of the pointer while taking motoring and power cards was eliminated by providing a means of putting pressure lines on the record. The modified indicator gives a complete record of the average cyclic variation in pressure per crank degree for any set of engine operating conditions which can be held constant for the period of time required to build up the composite card. The value of the record for accurate quantitative measurement is still questioned, although the maximum indicated pressure recorded on the motoring and power cards checks the readings of the balanced diaphragm type of maximum cylinder pressure indicator.
An analysis is made of the autogiro jump take-off, in which the kinetic energy of the rotor turning at excess speed is used to effect a vertical take-off. By the use of suitable approximations, the differential equation of motion of the rotor during this maneuver is reduced to a form that can be solved. Only the vertical jump was studied; the effect of a forward motion during the jump is discussed briefly. The results of model tests of the jump take-off have been incorporated in the paper and used to establish the relative accuracy of the results predicted from the analysis. Good agreement between calculation and experiment was obtained by making justifiable allowances.
This report presents a theoretical analysis of the lift on a trailing edge flap. An analytical expression has been derived which enables the computation of the flap load coefficient. The theoretical results seem to show a fair agreement with the meager experimental results which are available.
The von Karman-Millikan theory of laminar boundary layers presented in NACA Technical Report No. 504 is applied to the laminar boundary layer about an elliptic cylinder on which boundary-layer and pressure-distribution measurements were made. An outline of the procedure of the von Karman-Millikan method is given. Good agreement is obtained between the calculated and experimental results, indicating that the method may be applied generally to the laminar boundary layer about any body provided that an experimentally determined pressure distribution is available. It appears that for all Reynolds Number above 24,000 the separation point for the elliptic cylinder should occur at a constant distance behind the point of minimum pressure, provided that the boundary layer does not become turbulent.
A quantitative criterion of merit has been needed to assist airplane designers to incorporate satisfactory spinning characteristics into new designs. An approximate empirical criterion, based on the projected side area and the mass distribution of the airplane, has been formulated in a recent British report. In the present paper, the British results have been analyzed and applied to American designs. A simpler design criterion based solely on the type and the dimensions of the tail, has been developed: it is useful in a rapid estimation of whether a new design is likely to comply with the minimum requirements for safety in spinning.
A balanced diaphragm type of maximum cylinder pressure indicator was designed to give results consistent with engine operating conditions. The apparatus consists of a pressure element, a source of controlled high pressure and a neon lamp circuit. The pressure element, which is very compact, permits location of the diaphragm within 1/8 inch of the combustion chamber walls without water cooling. The neon lamp circuit used for indicating contact between the diaphragm and support facilitates the use of the apparatus with multicylinder engines.
This report presents test results obtained during an investigation of the performance of a single-cylinder, high-speed, compression-ignition test engine when using multiple-orifice fuel-injection valve nozzles in which the number and the direction of the orifices were varied independently.
Simple flight tests were made on ten conventional airplanes for the purpose of determining their action in the following two situations, which are generally thought to precede and lead to a large proportion of airplane crashes.
Curves of forced frequency against amplitude are presented for the conditions where the forced frequency is both increased and decreased into the resonant range. On the basis of these curves it is shown that the practical resonance frequency is the point where wrinkling first occurs and that the resonance frequency will be subject to considerable travel once permanent wrinkles appear in the vibrating shell. The decreasing mode of striking resonance is found to be by far the most destructive condition.
The paper beings with a brief discussion on the origin of the bending stresses in cantilever box beams under torsion. A critical survey of existing theory is followed by a summary of design formulas; this summary is based on the most complete solution published but omits all refinements considered unnecessary at the present state of development. Strain-gage tests made by NACA to obtained some experimental verification of the formulas are described next. Finally, the formulas are applied to a series of box beams previously static-tested by the U.S. Army Air Corps; the results show that the bending stresses due to torsion are responsible to a large extent for the free-edge type of failure frequently experienced in these tests.
Bending tests were made of two circular cylinders of corrugated aluminum-alloy sheet. In each test failure occurred by bending of the corrugations in a plane normal to the skin. It was found, after analysis of the effect of short end bays, that the computed stress on the extreme fiber of a corrugated cylinder is in excess of that for a flat panel of the same basic pattern and panel length tested as a pin-ended column. It is concluded that this increased strength was due to the effects of curvature of the pitch line. It is also concluded from the tests that light bulkheads closely spaced strengthen corrugated cylinders very materially.
Performance tests were made using a rectangular displacer arranged so that the combustion air was forced through equal passages at either end of the displacer into the vertical-disk combustion chamber of a single-cylinder, four-stroke-cycle compression-ignition test engine. After making tests to determine optimum displacer height, shape, and fuel-spray arrangement, engine-performance tests were made at 1,500 and 2,000 r.p.m. for a range of boost pressures from 0 to 20 inches of mercury and for maximum cylinder pressures up to 1,150 pounds per square inch. The engine operation for boosted conditions was very smooth, there being no combustion shock even at the highest maximum cylinder pressures. Indicated mean effective pressures of 240 pounds per square inch for fuel consumptions of 0.39 pound per horsepower-hour have been readily reproduced during routine testing at 2,000 r.p.m. at a boost pressure of 20 inches of mercury.
The results of take-off calculations are given for an application of simple trailing-edge flaps to two hypothetical flying boats, one having medium wing and power loading and consequently considerable excess of thrust over total resistance during the take-off run, the other having high wing and power loading and a very low excess thrust. For these seaplanes the effect of downward flap settings was: (1) to increase the total resistance below the stalling speed, (2) to decrease the get-away speed, (3) to improve the take-off performance of the seaplane having considerable excess thrust, and (4) to hinder the take-off of the seaplane having low excess thrust. It is indicated that flaps would allow a decrease in the high angles of wing setting necessary with most seaplanes, provided that the excess thrust is not too low.
In order to determine whether or not flaps could be expected to have any beneficial effect on take-off performance, the distances required to take off and climb to an altitude of 50 feet were calculated for hypothetical airplanes, corresponding to relatively high-speed types and equipped with several types of flap. The types considered are the Fowler wing, the Hall wing, the split flap, the balanced split flap, the plain flap, and the external-airfoil flap. The results indicate that substantial reductions in take-off distance are possible through the use of flaps, provided that the proper flap angle corresponding to a given set of conditions is used. The best flap angle for taking off varies inversely as power loading and, to a much smaller extent, varies inversely with wing loading. Apparently, the best take-off characteristics are provided by the type of device in which the flap forms an extension to the main wing as in the case of the Fowler wing and the external-airfoil flap.
A method is presented for calculating the aerodynamic forces on a moncylane wing, taking into account the elastic twisting of the wing due to these forces. The lift distribution along the span is calculated by the formulas of Amstutz as a function of the geometrical characteristics of the wing and of the twist at stations 60 and 90 percent of the semispan. The twist for a given lift distribution is calculated by means of influence lines. As a numerical example, the forces on a Swiss military D.2V airplane are calculated. Comparisons with the strip method and with the ordinary stress-analysis method are also given.
Tests on a Friez type cup anemometer have been made in the variable density wind tunnel of the Langley Memorial Aeronautical Laboratory to calibrate the instrument and to determine its suitability for velocity measurements of wind gusts. The instrument was calibrated against a Pitot-static tube placed directly above the anemometer at air densities corresponding to sea level, and to an altitude of approximately 6000 feet. Air-speed acceleration tests were made to determine the lag in the instrument reading. The calibration results indicate that there should be an altitude correction. It is concluded that the cup anemometer is too sluggish for velocity measurements of wind gusts.
Several improvements that have been made on commercially available carbon-monoxide indicators to make them more suitable for aircraft use are described. These improvements include an automatic flow regulator, which permits the use of a simplified instrument on aircraft where a source of suction is available, and a more reliable alarm attachment. A field method for testing instruments on standard samples of carbon monoxide is described. Performance data and instructions in operation and maintenance are given.
This report presents the results of tests made at a high value of the Reynolds Number in the N.A.C.A. variable-density wind tunnel to determine the aerodynamic characteristics of an airfoil as affected by fabric sag. Tests were made of two Gottingen 387 airfoils, one having the usual smooth surface and the other having a surface modified to simulate two types of fabric sag. The results of these tests indicate that the usual sagging of the wind covering between ribs has a very small effect on the aerodynamic characteristics of an airfoil.
According to Mr. L.D. Bell, of the Consolidated Aircraft Corporation, certain undesirable spinning characteristics of a commercial airplane were eliminated by the addition of a filler to the forward part of the wing to give it a sharp leading edge. To ascertain what aerodynamic effects result from such a change of section, two airfoils having sharp leading edges were tested in the variable-density wind tunnel. Both sections were derived by modifying the Gott. 398. The tests, which were made at a large value of the Reynolds Number, were carried to very large angles of attack to provide data for application to flight at angles of attack well beyond the stall. The characteristics of the sharp-nosed airfoils are compared with those of the normal Gott. 398 airfoil. Both of the sharp-nosed airfoils, which differ in the angle between the upper and lower surfaces at the leading edge, have about the same characteristics. As compared with the normal airfoil, the maximum lift is reduced by approximately 26 per cent, but the objectionable rapidly decreasing lift with angle of attack beyond the stall is eliminated; the profile drag of the section is slightly reduced in the range of the lift coefficient between 0.2 and 0.85, but at higher and lower lift coefficients the drag is increased.
Experimental measurements and theoretical calculations were made on an aircraft-type, single cylinder engine, in order to determine the physical nature of the inlet process, especially at high piston speeds. The engine was run at speeds from 1,500 to 2,600 r.p.m. (mean piston speeds of 1,370 to 2,380 feet per minute). Measurements were made of the cylinder pressure during the inlet stroke and of the power output and volumetric efficiency. Measurements were also made, with the engine not running, to determine the resistance and mass of air in the inlet valve port at various crank angles. Results of analysis indicate that mass has an appreciable effect, but friction plays the major part in restricting flow. The observed fact that the volumetric efficiency is considerably less than 100 percent is attributed to thermal effects. An estimate was made of the magnitude of these effects in the present case, and their general nature is discussed.
In this report charts are given showing the relation between time, velocities, and altitude for airplanes having various terminal velocities diving in a standard atmosphere. The range of starting altitudes is from 8,000 to 32,000 feet, and the terminal velocities vary from 150 to 550 miles per hour. A comparison is made between an experimental case and the results obtained from the charts. Examples pointing out the use of the charts are included.
Charts are presented for determining the performance of airplanes having variable-pitch propellers, the pitch of which is assumed to be adjusted to maintain constant speed for all rates of flight. The charts are based on the general performance equations developed by Oswald in reference 1, and are used in a similar manner. Examples applying the charts to airplanes having both supercharged and unsupercharged engines are included.
This report presents a convenient method for calculating the pitching-moment characteristics of tapered wings with sweepback and twist. The method is based on the fact that the pitching-moment characteristics of a wing may be specified by giving the value of the pitching moment at zero lift and the location of the axis about which the axis is constant. Data for calculating these characteristics are presented by curves which apply to wings having a linear distribution of twist along the span and which cover a large range of aspect ratios. The curves are given for wings having straight taper and distorted elliptical plan forms. The characteristics of wings of other shapes may be determined by interpolation.
The circular motion for airship-like bodies has thus far been calculated only for a prolate ellipsoid of revolution (reference 1, p.133 and reference 2). In this paper, however, the circular motion of elongated bodies of revolution more nearly resembling airships will be investigated. The results will give the effect of rotation on the pressure distribution and thus yield some information as to the stresses set up in an airship in circular flight.
Measurements were made of the circulation about the rectangular tip of a short-span airfoil passing through an artificial gust of known velocity gradient. A Clark Y airfoil of 30-centimeter chord was mounted on a whirling arm and moved at a velocity of 29 meters per second over a vertical gust with a velocity of nearly 7 meters per second. Flow angles were measured with a hot-wire apparatus. The rate at which the lift at the tips of a wing entering a gust is realized was found to be in satisfactory agreement with that predicted on the basis of the two-dimensional theory of von Karman and Sears.
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