ALBERT

Beschreibung: The relation between the elevator hinge moment parameters and the control forces for changes in forward speed and in maneuvers is shown for several values of static stability and elevator mass balance. The stability of the short period oscillations is shown as a series of boundaries giving the limits of the stable regions in terms of the elevator hinge moment parameters. The effects of static stability, elevator moment of inertia, elevator mass unbalance, and airplane density are also considered. Dynamic instability is likely to occur if there is mass unbalance of the elevator control system combined with a small restoring tendency (high aerodynamic balance). This instability can be prevented by a rearrangement of the unbalancing weights which, however, involves an increase of the amount of weight necessary. It can also be prevented by the addition of viscous friction to the elevator control system provided the airplane center of gravity is not behind a certain critical position. For high values of the density parameter, which correspond to high altitudes of flight, the addition of moderate amounts of viscous friction may be destabilizing even when the airplane is statically stable. In this case, increasing the viscous friction makes the oscillation stable again. The condition in which viscous friction causes dynamic instability of a statically stable airplane is limited to a definite range of hinge moment parameters. It is shown that, when viscous friction causes increasing oscillations, solid friction will produce steady oscillations having an amplitude proportional to the amount of friction.

Beschreibung: An investigation was made of the flow downstream from a "two-dimensional" grid formed of parallel rods. In both two and three dimensional jet fields there is a critical range of grid density below which the downstream flow is stable and above which it is unstable. The flow can be completely stabilized by means of an adequate lateral contraction beginning immediately after the grid or by use of a fine-mesh damping screen parallel to the grid plane and within a definite range of positions downstream from the grid.

Beschreibung: A study was made of the performance of a jet-propulsion system composed of an engine-driven blower, a combustion chamber, and a discharge nozzle. A simplified analysis is made of this system for the purpose of showing in concise form the effect of the important design variables and operating conditions on jet thrust, thrust horsepower, and fuel consumption. Curves are presented that permit a rapid evaluation of the performance of this system for a range of operating conditions. The performance for an illustrative case of a power plant of the type under consideration id discussed in detail. It is shown that for a given airplane velocity the jet thrust horsepower depends mainly on the blower power and the amount of fuel burned in the jet; the higher the thrust horsepower is for a given blower power, the higher the fuel consumption per thrust horsepower. Within limits the amount of air pumped has only a secondary effect on the thrust horsepower and efficiency. A lower limit on air flow for a given fuel flow occurs where the combustion-chamber temperature becomes excessive on the basis of the strength of the structure. As the air-flow rate is increased, an upper limit is reached where, for a given blower power, fuel-flow rate, and combustion-chamber size, further increase in air flow causes a decrease in power and efficiency. This decrease in power is caused by excessive velocity through the combustion chamber, attended by an excessive pressure drop caused by momentum changes occurring during combustion.

Beschreibung: 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.

Beschreibung: The effects of changes in aileron rigging between 2 deg up and 2 deg down on the stick forces were determined from wind-tunnel data for a finite-span wing model. These effects were investigated for ailerons deflecting equally in both directions and linearly with stick deflection. Data were analyzed for a Frise, a sealed internally balanced, and a beveled-trailing-edge aileron. The results of the analysis showed that only ailerons having linear hinge-moment characteristics are unaffected by changes in rigging and indicated that ailerons having decidedly nonlinear hinge-moment-coefficient curves, particularly for deflections near 0 deg, are very sensitive to changes in rigging.

Beschreibung: Pressure distribution measurements were made over an airfoil with slotted Frise aileron up to 0.76 Mach at various angles of attack and aileron defections. Section characteristics were determined from these pressure data. Results indicated loss of aileron rolling power for deflections ranging from -12 Degrees to -19 Degrees. High stick forces for non-differential deflections incurred at high speed, which were due to overbalancing tendency of up-moving aileron, may precipitate serious control difficulties. Detailed results are presented graphically.

Beschreibung: Methods are given of determining the potential flow plast an arbitrary cascade of airfoils and the inverse problem of determining an airfoil having a prescribed velocity distribution in cascade. Results indicated that Cartesian mapping function method may be satisfactorily extended to include cascades. Numerical calculation for computing cascades by Cartesian mapping function method is considerably greater than for single airfoils but much less than hitherto required for cascades. Detailed results are presented graphically.

Beschreibung: Flight tests were conducted on the OS2U-2 seaplane with simple circular-arc-type ailerons directly connected to the actuating torque tube. Two aileron test installations were made, differing only in the inclination of the projecting surface with the wing's upper surface. The lateral-control characteristics of the airplane were determined from data obtained in stalls and rudder-fixed aileron rolls. The revised ailerons were deficient in maximum rolling effectiveness, but were capable of controlling the rolling tendencies of the airplane near the stall.

Beschreibung: The effects of jet-motor operation on the stability and control characteristics of two fighter-type airplanes as determined by wind-tunnel tests of 1/5-scale models are presented. It is shown that the action of the jets is to cause a small loss in stick-fixed stability which is predictable from known theories.

Beschreibung: Tank tests were made of a hull model of the Hughes-Kaiser cargo airplane for estimates of take-off performance and maximum gross load for take-off. At hump speeds, with the model free to trim, the trim and resistance were high, which resulted in a load-resistance ratio of approximately 4.0 for a gross load coefficient of 0.75. With a 4000,000-lb load, the full size craft may take off in 69 sec over a distance of 5600 ft.

Beschreibung: No abstract available

Beschreibung: Results of flight tests indicate that profile-drag coefficients which were obtained with the low-drag airfoils were lower than with the conventional types over the range of light coefficients tested. For comparable conditions of the lift coefficient and Reynolds Number, the low-drag airfoils have profile-drag coefficients which may be 27 percent lower than the profile drag of the conventional airfoils tested. Detailed results are presented graphically.

Beschreibung: Rough conventional, smooth conventional, and laminar-flow or low-drag sections were tested. The items covered are rotor thrust for fixed power in hovering, range and endurance at cruising speed, and power required at high-forward speed. Calculations indicated that a smooth conventional section gives marked performance gains. Smaller gains are obtainable by using a low-drag section. At high speeds or loads the low-drag section is inferior to the smooth conventional section.

Beschreibung: An investigation was made of the cooling characteristics of a P and W R-2800 engine with NACA short-nose high inlet-velocity cowling. The internal aerodynamics of the cowling were studied for ranges of propeller-advance ratio and inlet-velocity ratio obtained by deflection of cowling flaps. Tests included variations of engine power, fuel/air ratio and cooling-air pressure drop. Engine cooling data are presented in the form of cooling correlation curves, and an example for calculation of cooling requirements in flight is included.

Beschreibung: No abstract available

Beschreibung: In open box beams subjected to torsion, secondary stresses arise owing to lateral bending of the spar caps. The present paper outlines a simple method for estimating the magnitude of these stresses and gives the results of tests of an open box beam in the neighborhood of a discontinuity where the cover changed from the top to the bottom of the box.

Beschreibung: Results are presented for tests of two wings, an NACA 230-series wing and a highly-cambered NACA 66-series wing on a twin-engine pursuit airplane. Auxiliary control flaps were tested in combinations with each wing. Data showing comparison of high-speed aerodynamic characteristics of the model when equipped with each wing, the effect of the auxiliary control flaps on aerodynamic characteristics, and elevator effectiveness for the model with the 66-series wing are presented. High-speed aerodynamic characteristics of the model were improved with the 66-series wing.

Beschreibung: A correlation of what are believed to be the most reliable data available on duct components of aircraft power-plant installations is presented. The information is given in a convenient form and is offered as an aid in designing duct systems and, subject to certain qualifications, as a guide in estimating their performance. The design and performance data include those for straight ducts; simple bends of square, circular, and elliptical cross sections; compound bends; diverging and converging bends; vaned bends; diffusers; branch ducts; internal inlets; and an angular placement of heat exchangers. Examples are included to illustrate methods of applying these data in analyzing duct systems. (author)

Beschreibung: Correlation is established between aerodynamic characteristics of control surfaces in two-dimensional and three-dimensional flow. Slope of lift curve was affected little by overhang and balance-nose shape, but increased by sealing flap-nose gap. Effectiveness of balancing tab was same for sealed plain flap and unsealed overhang flap. Changes in hinge-moment coefficient were diminished by sealing gap. Values measured by three-dimensional flow disagreed with two-dimensional flow values until aspect ratio corrections were made.

Beschreibung: An investigation was made in the LMAL 7- by 10-foot wind tunnel of a NACA 23021 airfoil with a double slotted flap having a chord 32 percent of the airfoil chord (0.32c) to determine the aerodynamic section characteristics with the flaps deflected at various positions. The effects of moving the fore flap and rear flap as a unit and of deflecting or removing the lower lip of the slot were also determined. Three positions were selected for the fore flap and at each position the maximum lift of the airfoil was obtained with the rear flap at the maximum deflection used at that fore-flap position. The section lift of the airfoil increased as the fore flap was extended and maximum lift was obtained with the fore flap deflected 30 deg in the most extended position. This arrangement provided a maximum section lift coefficient of 3.31, which was higher than the value obtained with either a 0.2566c or a 0.40c single-slotted-flap arrangement and 0.25 less than the value obtained with a 0.4c double-slotted-flap arrangement on the same airfoil. The values of the profile-drag coefficient obtained with the 0.32c double slotted flap were larger than those for the 0.2566c or 0.40c single slotted flaps for section lift coefficients between 1.0 and approximately 2.7. At all values of the section lift coefficient above 1.0, the 0.40c double slotted flap had a lower profile drag than the 0.32c double slotted flap. At various values of the maximum section lift coefficient produced by various flap defections, the 0.32c double slotted flap gave negative section pitching-moment coefficients that were higher than those of other slotted flaps on the same airfoil. The 0.32c double slotted flap gave approximately the same maximum section lift coefficient as, but higher profile-drag coefficients over the entire lift range than, a similar arrangement of a 0.30c double slotted flap on an NACA 23012 airfoil.

Beschreibung: Results of subject tests indicate the difficulty of obtaining closely balanced rudder surfaces for most tail assemblies with shielded horns and maintaining a near zero rate-of-change of hinge-moment coefficient without an additional balancing device. A comparison is made between shielded and unshielded horn test results. Pressure distribution and tuft tests of flow over different shaped horns showed higher critical speed for medium-taper nosed horn. The trim tab nose shape had little effect on tab test results.

Beschreibung: Wind-tunnel tests, investigating low drag wing performance in small-scale tests, showed a large increase in minimum drag coefficient, and a decrease of maximum lift coefficient occurred with decreasing Reynolds Number above certain designated values. The lift-curve slope varied up to 6% between high and low turbulence levels. Low Reynolds Number test data are unreliable for low drag airfoils either to estimate full-scale characteristics or to determine merits of airfoils for higher Reynolds numbers.

Beschreibung: An investigation has been made in the Langley free-flight tunnel to obtain an experimental verification of the theoretical rudder-free stability characteristics of an airplane model equipped with conventional rudders having negative floating tendencies and negligible friction. The model used in the tests was equipped with a conventional single vertical tail having rudder area 40 percent of the vertical tail area. The model was tested both in free flight and mounted on a strut that allowed freedom only in yaw. Tests were made with three different amounts of rudder aerodynamic balance and with various values of mass, moment of inertia, and center-of-gravity location of the rudder. Most of the stability derivatives required for the theoretical calculations were determined from forced and free-oscillation tests of the particular model tested. The theoretical analysis showed that the rudder-free motions of an airplane consist largely of two oscillatory modes - a long-period oscillation somewhat similar to the normal rudder-fixed oscillation and a short-period oscillation introduced only when the rudder is set free. It was found possible in the tests to create lateral instability of the rudder-free short-period mode by large values of rudder mass parameters even though the rudder-fixed condition was highly stable. The results of the tests and calculation indicated that for most present-day airplanes having rudders of negative floating tendency, the rudder-free stability characteristics may be examined by simply considering the dynamic lateral stability using the value of the directional-stability parameter Cn(sub p) for the rudder-free condition in the conventional controls-fixed lateral-stability equations. For very large airplanes having relatively high values of the rudder mass parameters with respect to the rudder aerodynamic parameters, however, analysis of the rudder-free stability should be made with the complete equations of motion. Good agreement between calculated and measured rudder-free stability characteristics was obtained by use of the general rudder-free stability theory, in which four degrees of lateral freedom are considered. When this assumption is made that the rolling motions alone or the lateral and rolling motions may be neglected in the calculations of rudder-free stability, it is possible to predict satisfactorily the characteristics of the long-period (Dutch roll type) rudder-free oscillation for airplanes only when the effective-dihedral angle is small. With these simplifying assumptions, however, satisfactory prediction of the short-period oscillation may be obtained for any dihedral. Further simplification of the theory based on the assumption that the rudder moment of inertia might be disregarded was found to be invalid because this assumption made it impossible to calculate the characteristics of the short-period oscillations.

Beschreibung: Tests were conducted at dynamic pressure of 50 lb per square foot with lift drag and pitch moment measurements throughout useful angle of attack range for constant flap deflection and position of a low-drag airfoil. Two slots were investigated and practical flap paths were selected for each Slot shape had a negligible effect on the maximum lift coefficient flap deflected, the rounded-entry slot had lower profile drag.

Beschreibung: Two graphical methods are presented for determining the stick-free neutral point, and they are extensions of the methods commonly used to determine the stick-free neutral point. A mathematical formula for computing the stick-free neutral point is also given. These methods may be applied to determine approximately the increase in tail size necessary to shift the neutral point (stick fixed or free) to any desired location on an airplane having inadequate longitudinal stability.

Beschreibung: No abstract available

Beschreibung: The effect of several armament installations on the drag of a 1/8 scale model of the B-32 airplane was determined. Turrets in the following positions were tested: nose, tail, upper forward, upper aft, and lower. The nose and tail turrets were each equipped with two.50-caliber guns. Upper turrets were of three types: two.50-caliber guns, four.50-caliber guns, and 20-millimeter cannon. Lower turrets were of two types: two.50-caliber guns and four.50-caliber guns. The effect of streamlining the upper two- and four-gun turrets and of extending the lower two-gun turret was determined. The tests were conducted in the Langley 19-foot. pressure tunnel at a Reynolds number of approximately 2,960,000 and a Mach number of 0.13. Large increases in drag coefficient were caused by the complete armament installations. At a lift coefficient of 0. 4 the installations with nonstreamlined upper turrets and the lower turret retracted increased the drag coefficient by 0.0022 and 0.0027 for the two-gun and four-gun turret installations, respectively. Streamlining the upper turrets reduced the drag of these installations by approximately 40 percent with the upper turrets streamlined, the drag increase was about the same for either the two- or four-gun turret installation. The streamlined two-cannon upper turrets increased the drag about the same amount as the two-gun upper turrets that were not streamlined. Extension of the lower turret. increased the drag slightly more than the whole streamlined gun-turret installation.

Beschreibung: This report gives the results of tests on a rectangular wing model with a 20% full spun split flap, conducted on the whirling arm at the Daniel Guggenheim Airship Institute in Akron, Ohio. The effect of a ground board on the lift and pitching moment was measured. The ground board consisted of an inclined ramp rising up in the test channel to a level floor extending for some distance parallel to the model path. The path of the wing model with respect to the ground board accordingly represented with comparative exactness an airplane coming in for a landing. The ground clearances over the level portion of the board varied from 0 6 to 1,6 chord lengths. Results are given in the standard dimensionless coefficients plotted versus angle of attack for a particular ground clearance. The effect of the ground board is to increase the lift coefficient for a given angle of attack all the way up the stall. The magnitude of the increase varies both with the ground clearance and the angle of attack. The effect on the pitching moment coefficient is not so readily apparent due to experimental difficulties but, in general, the diving moment increases over the ground board. This effect is apparent principally at the high angles of attack. An exception to this effect occurs with flaps deflected at the lowest ground clearance (0.6 chords). Here the diving moment decreases over the ground board.

Beschreibung: In this report a method is presented for the calculation of the profile drag of airfoil sections. The method requlres only a knowledge of the theoretical velocity distribution and can be applied readily once this dlstribution is ascertained. Comparison of calculated and experimental drag characteristics for several airfoils shows a satisfactory agreement. Sample calculatlons are included.

Beschreibung: This report outlines the flight conditions that are usually critical in determining the design of components of an airplane which affect its stability and control characteristics. The wind-tunnel tests necessary to determine the pertinent data for these conditions are indicated, and the methods of computation used to translate these data into characteristics which define the flying qualities of the airplane are illustrated.

Beschreibung: Problems relating to the stability and control of tailless airplanes are discussed in consideration of contemporary experience and practice.

Beschreibung: Field measurements were made on a fighter airplane to determine the approximate magnitude of the horizontal tail loads in accelerated flight. In these flight measurements, pressures at a few points were used as an index of the tail loads by correlating these pressures with complete pressure-distribution data obtained in the NACA full-scale tunnel. In addition, strain gages and motion pictures of tail deflections were used to explore the general nature and order of magnitude of fluctuating tail loads in accelerated stalls.

Beschreibung: An experimental investigation concerned primarily with the extension of test data on the drag of revolving disks, cylinders, and streamline rods to high Mach numbers and Reynolds numbers is presented.

Beschreibung: The relation between the elevator hinge-moment parameters and the control-forces for changes in forward speed and in maneuvers is shown for several values of static stability and elevator mass balance.

Beschreibung: In many studies, especially of nonstationary flight motion, it is necessary to determine the angular velocities at which the airplane rotates about its various axes. The three-component recorder is designed to serve this purpose. If the angular velocity for one flight attitude is known, other important quantities can be derived from its time rate of change, such as the angular acceleration by differentiations, or - by integration - the angles of position of the airplane - that is, the angles formed by the airplane axes with the axis direction presented at the instant of the beginning of the motion that is to be investigated.

Beschreibung: The wall interference is obtained for a wind tunnel of elliptic section for the two cases of closed and open working sections. The approximate and exact methods used gave results in practically good agreement. Corresponding to the result given by Glauert for the case of the closed rectangular section, the interference is found to be a minimum for a ratio of minor to major axis of 1:square root of 6 This, however, is true only for the case where the span of the airfoil is small in comparison with the width of the tunnel. For a longer airfoil the favorable ellipse is flatter. In the case of the open working section the circular shape gives the minimum interference.

Beschreibung: The present report deals with the determination of the impact stresses in the bulkhead floors of a seaplane bottom. The dynamic problem is solved on the assumption of a certain elastic system, the floor being assumed as a weightless elastic beam with concentrated masses at the ends (due to the mass of the float) and with a spring which replaces the elastic action of the keel in the center. The distributed load on the floor is that due to the hydrodynamic force acting over a certain portion of the bottom. The pressure distribution over the width of the float is assumed to follow the Wagner law. The formulas given for the maximum bending moment are derived on the assumption that the keel is relatively elastic, in which case it can be shown that at each instant of time the maximum bending moment is at the point of juncture of the floor with the keel. The bending moment at this point is a function of the half width of the wetted surface c and reaches its maximum value when c is approximately equal to b/2 where b is the half width of the float. In general, however, for computing the bending moment the values of the bending moment at the keel for certain values of c are determined and a curve is drawn. The illustrative sample computation gave for the stresses a result approximately equal to that obtained by the conventional factory computation.

Beschreibung: The wide use of diffusers, in various fields of technology, has resulted in several experimental projects to study the action and design of diffusers. Most of the projects dealt with steam (steam turbine nozzles). But diffusers have other applications - that is, ventilators, smoke ducts, air coolers, refrigeration, drying, and so forth. At present there is another application for diffusers in wind-tunnel design. Because of higher requirements and increased power of such installations more attention must be paid to the correctness of work and the decrease in losses due to every section of the tunnel. A diffuser, being one of the component parts of a tunnel , can in the event of faulty construction introduce considerable losses. Therefore, in the design of the new CAHI wind tunnel, it was suggested that an experimental study of diffusers be made, with a view to applying the results to wind tunnels. The experiments conducted by K. K. Baulin in the laboratories of CAHI upon models of diffusers of different cross sections, lengths, and angles of divergence, were a valuable source of experimental data. They were of no help, however, in reaching any conclusion regarding the optimum shape because of the complexity and diversity of the factors which all appeared simultaneously, thereby precluding the.study of the effects of any one factor separately. On the suggestion of the director of the CAHI,Prof. B. N. Ureff, it was decided to experiment on a two-dimensional diffuser model and determine the effect, of the angle of divergence. The author is acquainted with two experimental projects of like nature: the first was conducted with water, the other with air. The first of these works, although containing a wealth of experimental data, does not indicate the nature of flow or its relation to the angle of divergence. The second work is limited to four angles - that is, 12 deg, 24 deg, 45 deg, 90 deg. The study of this diffuser did not supply any information about the effect of smaller angles which, because of their advantages, are more commonly used, The author was able to acquaint himself with the second work only after the experiments were started. For these reasons, as well as because on the basis of those works no conclusion can be reached regarding the nature of flow distribution, of eddies, and so forth, experimental work was continued. The need for determining flow patterns follows from the fact that from them are determined methods of measurement - that is, the determination of velocities by means of the pitot tube, which, as is well known, gives correct indications only when placed with its axis parallel to the axis of flow. The data contained in this report were obtained from experiments conducted by the Aerodynamical Laboratories of the CAHI. The solutions to some. of the mathematical problems connected with the experiments are due to Prof. S. A, Chapligin.

Beschreibung: Detail calculations are presented of the shifts in stick-fixed neutral point of the Republic XF-12 airplane due to the windmilling propellers and to the fuselage. The results of these calculations differ somewhat from those previously made for this airplane by Republic Aviation Corporation personnel under the direction of Langley flight division personnel. Due to these differences the neutral point for the airplane is predicted to be 37.8 percent mean aerodynamic chord, instead of 40.8 percent mean aerodynamic chord as previously reported.

Beschreibung: Pressure distribution and spray measurements were carried out on rectangular flat and V-bottom planing surfaces. Lift, resistance, and center of pressure data are analyzed and it is shown how these values may be computed for the pure planing procees of a flat or V-bottom suface of arbitrary beam, load and speed, the method being illustrated with the aid of an example.

Beschreibung: An analysis is made of the stability of an airplane with ailerons free, with particular attention to the motions when the ailerons have a tendency to float against the wind. The present analysis supersedes the aileron investigation contained in NACA Report No. 709. The equations of motion are first written to include yawing and sideslipping, and it is demonstrated that the principal effects of freeing the ailerons can be determined without regard to these motions. If the ailerons tend to float against the wind and have a high degree of aerodynamic balance, rolling oscillations, in addition to the normal lateral oscillations, are likely to occur. On the basis of the equations including only the rolling motion and the aileron deflection, formulas are derived for the stability and damping of the rolling oscillations in terms of the hinge moment derivatives and other characteristics of the ailerons and airplane. Charts are also presented showing the oscillatory regions and stability boundaries for a fictitious airplane of conventional proportions. The effects of friction in the control system are investigated and discussed. If the ailerons tend to trail with the wind, the condition for stable variation of stick force with aileron deflection is found to determine the amount of aerodynamic balance that may be used. If the ailerons tend to float against the wind, the period and damping of the rolling oscillations are found to be satisfactory (in a mass-balanced system) so long as the restoring moment is not completely balanced out. Unbalanced mass behind the hinge, however, has an unfavorable effect on the damping of the oscillations and so shifts the boundary that close aerodynamic balance may not be attainable.