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National Advisory Committee for Aeronautics. Research Abstracts NOVEMBER 17T 1953 NO. 52 CURRENT NACA REPORTS NACA Rept. 1094 AN EXPERIMENTAL INVESTIGATION OF TRAN SONIC FLOW PAST TWODIMENSIONAL WEDGE AND CIRCULARARC SECTIONS USING A MACH ZEHNDER INTERFEROMETER. Arthur Earl Bryson, Jr., California Institute of Technology. 1952. ii, 33p. diagrs., photos. (NACA Rept. 1094. Formerly TN 2560) Interferometer measurements are given of the flow fields near twodimensional wedge and circulararc sections at zero angle of attack. Pressure distribu tions and drag coefficients as functions of Mach num ber were obtained and the wedge data are compared with theory. It is shown that the local Mach number at any point on the surface of a finite three dimen sional body or an unswept twodimensional body, moving through an infinite fluid, has a stationary value at Mach number 1 and, in fact, remains nearly constant for a range of speeds below and above Mach number 1. On the basis of this concept and the experimental data, pressure distributions and drag coefficients for the wedge and circulararc sections are presented throughout the entire transonic range of velocities. NACA Rept. 1097 STRESSES IN A TWOBAY NONCIRCULAR CYLIN DER UNDER TRANSVERSE LOADS. George E. Griffith. 1952. ii, 12p. diagrs.. 3 tabs. (NACA Rept. 1097. Formerly TN 2512) A method, taking into account the effects of flexibility and based on a general eighthorder differential equation, is presented for finding the stresses in a twobay, noncircular cylinder the cross section of which can be composed of circular arcs. Numerical examples are given for two cases of ring flexibility for a cylinder of doubly symmetrical (essentially el liptic) cross section, subjected to concentrated radial, moment, and tangential loads. The results parallel those already obtained for shells with circu lar rings. NACA Rept. 1112 HYDROCARBON AND NONHYDROCARBON DERIV ATIVES OF CYCLOPROPANE. Vernon A. Slabey, Paul H. Wise and Louis C. Gibbons. 1953. ii, 18p. diagram 4 tabs. (NACA Rept. 1112) The methods used to prepare and purify 19 hydro carbon derivatives of cyclopropane are discussed. Of these hydrocarbons, 13 were synthesized for the first time. In addition to the hydrocarbons, six cyclopropylcarbinols, five alkyl cyclopropyl ketones, three cyclopropyl chlorides, and one cyclopro panedicarboxylate were prepared as synthesis inter mediates. The melting points, boiling points, re fractive indices, densities, and, in some instances, heats of combustion of both the hydrocarbon and nonhydrocarbon derivatives of cyclopropane were determined. These data and the infrared spectrum of each of the 34 cyclopropane compounds are pre sented herein. The infrared absorption bands characteristic of the cyclopropyl ring are discussed, and some observations are made on the contribution of the cyclopropyl ring to the molecular refractions of cyclopropane compounds. NACA Rept. 1113 sriCTRUM OF TURBULENCE IN A CONTRACTING STREAM. H. S. Ribner and M. Tucker. 1953. ii, 17p. diagrs., tab. (NACA Rept. 1113. Formerly TN 2606) Spectrum concepts are employed to study the selec tive effect of a stream contraction on longitudinal and lateral turbulent velocity fluctuations. By consider ation of the effect of stream contraction on a single plane wave, the effect on spectrum and correlation tensors of the turbulence is determined. Weak turbu lence and an inviscid fluid are postulated; compress ibility of mean flow only is taken into account. For axisymmetric contraction and isotropic initial turbu lence, explicit results s'_ eebaLned. The one dimen sional longitudinal sf'trum 'ffitd to be markedly distorted. The oa'eqhive efll re i t reaction on lon gitudinal and lapra components of turbulence is found to be given unique regardless a of the iso tropic spectrum; pomparisn 3 !peri ent ismade. NACA TN 3022 METHOD FOR STUDYING HEyw6P'ER LONGI TUDINAL MANEUVER STATITLITY.' Kenneth B. Amer. October 1953. f, 52p. *diagrs., photos., 2 tabs. (NACA TN 3022) A theoretical analysis of helicopter maneuver stabil ity is made and the results are compared with ex perimental results for both a single and a tandem rotor helicopter. Techniques are described for measuring in flight the significant stability deriva tives for use with the theory to aid in design studies of means for achieving marginal maneuver stability for a prototype helicopter. NACA TN 3026 ELECTROSTATIC SPARK IGNITIONSOURCE HAZARD IN AIRPLANE CRASHES. Arthur M. Busch. October 1953. 28p. diagrs., photos., tab. (NACA TN 3026) The hazard of igniting airplane crash fires by elec * AVAILABLE ON LOAN ONLY ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON 2s, D.C., CITING CODE NUMBER ABOVE EACH TITLE, THE REPORT TITLE AND AUTHOR. eg.9, /Jo0? 2 trostatic sparks, generated when detached airplane parts fly through clouds of dust and fuel mist, was investigated. Within the limits of variables studied, the rates with which airplane wreckage collected a charge were directly proportional to the rate that clay dust or fuel mist was intercepted. Maximum rates of experimental electrification were used to relate energy accumulation to wreckage sizes and trajectories and to estimate minimum hazardous wreckage sizes and trajectories. Comparison of sizes and trajectories of wreckage shown in motion pictures of airplane crashes with these estimated sizes and trajectories indicated that the hazard is small. Of the remedial measures considered, poly ethylene coatings were found to offer promise of protection against electrostatic spark ignition. NACA TN 3027 INFLUENCE OF ROTORENGINE TORSIONAL OSCILLATION ON CONTROL OF GASTURBINE ENGINE GEARED TO HELICOPTER ROTORS. John C. Sanders. October 1953. 40p. diagrs., photos., 2 tabs. (NACA TN 3027) Equations were developed for the torsional motion of a gasturbine engine geared to a helicopter rotor in which the rotor blades were hinged to the rotor shaft. The rotor system was simplified to yield simple thirdorder equations that can be used in the analysis of engine control. Comparison of the system re sponse calculated from these equations with the ex perimentally observed frequency response of a rotor from a 2500pound helicopter showed satisfactory agreement. Calculations showed that the torsional motion arising from the hinged construction of the blades contributed to instability of the enginespeed and enginetorque controls. Trends in stability and response of controls with increasing weight of heli copters were investigated. NACA TN 3028 THE COMPRESSIBLE LAMINAR BOUNDARY LAYER WITH HEAT TRANSFER AND SMALL PRESSURE GRADIENT. George M. Low. APPENDIX B. NUMERICAL SOLUTION OF DIFFER ENTIAL EQUATIONS. Lynn U. Albers. October 1953. 68p. diagrs., 7 tabs. (NACA TN 3028) A perturbation method for the calculation of velocity and temperature profiles, skinfriction and heat transfer characteristics for twodimensional com pressible laminar boundary layers with heat transfer and small arbitrary gradient is presented. The permissible pressure gradients include those of a form and magnitude usually encountered over slender aerodynamic shapes in supersonic flight. The combined effects of heat transfer and pressure gradi ent on boundarylayer characteristics are demon strated by applying the results of the analysis to two representative wings. NACA TN 3032 AN ANALYTICAL STUDY OF THE EFFECT OF AIR PLANE WAKE ON THE LATERAL DISPERSION OF AERIAL SPRAYS. Wilmer H. Reed, III. October 1953. 46p. diagrs., 3 tabs. (NACA TN 3032) An analysis is made to determine the trajectories and deposit of aerial spray droplets which are issued into the air disturbances generated by an agricultural NACA RESEARCH ABSTRACTS NO. 52 airplane. Various nozzle arrangements and droplet size spectra are considered with a view to improving the uniformity and effective width of the deposUt. 4 NACA TN 3034 GRAPHICAL SOLUTION OF SOME AUTOMATIC CONTROL PROBLEMS INVOLVING SATURATION EFFECTS WITH APPLICATION TO YAW DAMPERS FOR AIRCRAFT. William H. Phillips. October 1953. 41p. diagrs. (NACA TN 3034) A graphical method is presented for determining the motion of a freely oscillating system of one degree of freedom stabilized by a controlling device which applies control force in proportion to the displace ment of the system, to its rate of change of dis placement, or both. The controlling member is assumed to have limitations on its maximum deflec  tion and on its maximum rate of movement. Several examples concerned with the application of yaw dampers to aircraft are presented to illustrate the method, and some conclusions regarding the varia tion of the stability of the motion with amplitude are obtained. NACA TN 3035 A PRELIMINARY STUDY OF THE PROBLEM OF DESIGNING HIGHSPEED AIRPLANES WITH SATIS FACTORY INHERENT DAMPING OF THE DUTCH ROLL OSCILLATION. John P. Campbell and Marion 0. McKinney, Jr. October 1953. 40p. diagrs., 4 tabs. (NACA TN 3035) This paper presents some preliminary results of a theoretical investigation to determine what design features appear most important in designing modern highperformance airplanes to have the greatest possible inherent stability of the Dutch roll oscilla tion in order that the need for artificial stabilizing devices can be minimized. These preliminary results cover the case of fighter airplanes at sub sonic speeds. NACA TN 3036 THE FLOW ABOUT A SECTION OF A FINITE ASPECTRATIO NACA 0015 AIRFOIL ON A TRAN SONIC BUMP. Jack A. Mellenthin. October 1953. 30p. diagrs., photos., tab. (NACA TN 3036) Pressure distributions were measured at one span wise station on a semispan rectangular wing model having an NACA 0015 airfoil section and a moderate aspect ratio. The tests were conducted on a tran sonic bump at Mach numbers from 0.4 to 1.06. Pressuredistribution plots were integrated to obtain the section lift, drag, and pitchingmoment coeffi cients. At a fixed angle, a region developed over the airfoil surface wherein the local Mach number remained nearly constant as the freestream Mach number was increased above the critical value. This region covered essentially the whole chord of the air foil at freestream Mach numbers near unity. NACA TN 3037 COUNTING METHODS AND EQUIPMENT FOR MEANVALUE MEASUREMENTS IN TURBULENCE RESEARCH. H. W. Liepmann and M. S. Robinson, California Institute of Technology. October 1953. 49p. diagrs., photos. (NACA TN 3037) NACA RESEARCH ABSTRACTS NO.52 Methods of measuring the probability distributions and mean values of random functions as encountered in turbulence research were studied. Applications to the measurement of probability distributions of the axial velocity fluctuation u(t) and its derivative du dt in isotropic turbulence are shown. The assumption of independent probabilities of u(t) and du dt was investigated and the results indicate that the assumption is satisfied within a few percent and that there is no evidence that the systematic differ ence between the microscale of turbulence measured from zero counts and measured independently can be traced entirely to the statistical dependence of u and du dt. The chronological development of apparatus is described, concluding with the 10 channel statistical analyzer. NACA TN 3042 HIGHFREQUENCY PRESSURE INDICATORS FOR AERODYNAMIC PROBLEMS. Y. T. Li, Massachusetts Institute of Technology. November 1953. 52p. diagrs., photos., 4 tabs. (NACA TN 3042) Three different types of pressure indicators are discussed. Each of these indicators has a unique feature, but all are designed with an attempt to combine both highfrequency response and high resolving power into one instrument. Of the mechanical electricaltransducer type of pressure indicator, the wire strain gage leads in simplicity. The capacitance type is more versatile because it permits the use of very high frequency carrier sys tems and thereby cuts down the effective interference in the electronic system. The system utilizing the stretching of a bariumtitanate disk produces large signals and results in compact design, but it can only be used for dynamic measurements when tem perature variations are slight. Five different types of pressure receivers were tested. The flat diaphragm type leads the others in simplicity, the sphericaldiaphragm type exceeds in dynamic per formance, and the catenarydiaphragm type is the one least affected by temperature change. NACA TN 3045 ANALOGY BETWEEN MASS AND HEAT TRANSFER WITH TURBULENT FLOW. Edmund E. Callaghan. October 1953. 19p. diagrs. (NACA TN 3045) An analysis of combined heat and mass transfer from a flat plate has been made in terms of Prandtl's simplified physical concept of the turbulent boundary layer. The results of the analysis show that for con ditions of reasonably small heat and mass transfer, the ratio of the mass and heattransfer coefficients is dependent on the Reynolds number of the boundary layer, the Prandtl number of the medium of diffusion, and the Schmidt number of the diffusing fluid in the medium of diffusion. For the particular case of water evaporating into air, the ratio of mass transfer coefficient to heattransfer coefficient is found to be slightly greater than unity. NACA TM 1360 CONCERNING THE FLOW ABOUT RINGSHAPED COWLINGS. PART XII. TWO NEW CLASSES OF CIRCULAR COWLS. (Uber die Stromung an 3 ringtfrmigen Verkleidungen. XII Mitteilung: Zwei neue Klassen von Ringhaubenj. Dietrich Kiichemann and Johanna Weber. October 1953. 72p. diagrs., 3 tabs. (NACA TM 1360. Trans. from Zentrale fir wissenschaftliches Berichtswesen der Luftfahrtforschung, Berlin. UM 3111) For application in practice for annular radiator fairings and similar arrangements, two new classes of circular cowls are developed by theoretical method, and investigated in a systematic test series regarding their behavior under various working conditions. NACA RM E53G03 EFFECTIVE THERMAL CONDUCTIVITIES OF MAGNESIUM OXIDE, STAINLESS STEEL, AND URANIUM OXIDE POWDERS IN VARIOUS GASES. C. S. Elan and R. G. Deissler. October 1953. 18p. diagrs., photos., tab. (NACA RM E53G03) As a part of a general investigation of the effective thermal conductivities of powders, tests were con ducted to determine the conductivity of magnesium oxide, stainless steel, and uranium oxide powders in various gases at temperatures between 1200 and 14550 F. Fair agreement was obtained between con ductivities calculated from experimental data for fine magnesium oxide and stainless steel powders and those calculated from a simplified analysis from a previous investigation, although the experimental values are somewhat higher. Runs were also made to determine the effect of gas pressure on effective thermal conductivity. NACA RM E53H31 MINIMUM SPARKIGNITION ENERGIES OF 12 PURE FUELS AT ATMOSPHERIC AND REDUCED PRESSURE. Allen J. Metzler. October 1953. 28p. diagrs., 5 tabs. (NACA RM E53H31) Minimum sparkignition energies for 12 pure fuels were measured at reduced pressure, and the data ob tained were extrapolated to 1 atmosphere. The fuels investigated included normal and cycloparaffins, olefins, carbon disulfide, and oxygenated compounds such as an alcohol, ether, propylene oxide, and tetrahydropyran; these fuels were ignited at reduced pressures by capacitance sparks of controlled dura tion. The minimum ignition energies obtained are related to the pressure, the quenching distance, and the maximum fundamental flame velocity of the fuel air mixture. Also, the experimental data obtained are applied to two correlations of sparkigrunition energies to check the data of this investigation with that of others. NACA RM L53I118a FACTORS AFFECTING TRANSITION AT SUPER SONIC SPEEDS. K. R. Czarnecki and Archibald R. Sinclair. November 1953. 13p. diagrs. (NACA RM L53Il8a) The paper surveys the available material and summa rizes what is known to date about boundarylayer transition at supersonic speeds. Variables studied include Mach number, Reynolds number, pressure gradients, heat transfer, surface roughness, and angle of attack. The discussion is limited to bodies of revolution because similar reliable data for wings is lacking. BRITISH REPORTS N25356* Aeronautical Research Council (Gt. Brit.) WINGFUSELAGE FLUTTER OF LARGE AERO PLANES. W. P. Jones. 1953. 46p. diagrs., 7 tabs. (ARC R & M 2656. Formerly ARC 11,024; 0.688) A general theoretical method is described which takes into account a large number of degrees of free dom and is based on the design data for the airplane. The problem specifically investigated is the symmet rical flutter of a particular aircraft. Twelve degrees of freedom are assumed to cover pitching and trans lational motion of the whole airplane, flexure and torsion of the wings, and fuselage vertical bending. The tailplane is regarded as rigid. In the case con sidered, estimates indicate that the lowest critical speed is well above the maximum design speed of the airplane. The influence of the additional degrees of freedom associated with movements of the control surfaces is not considered. N26658* Royal Aircraft Establishment (Gt. Brit.) CALCULATIONS OF THE PRESSURE DISTRIBU TIONS AND BOUNDARY LAYER DEVELOPMENT ON A BODY OF REVOLUTION WITH VARIOUS PARA BOLIC AFTERBODIES AT SUPERSONIC SPEEDS. L. E. Fraenkel. February 1953. 56p. diagrs., 4 tabs. (RAE Aero 2482) Detailed calculations are made of the flow over a series of bodies at Mach numbers of 1.2, 1.4, and 1.6 and Reynolds numbers of 48 to 72 millions. The bodies consist of a basic forebody and parallel portion to which are added truncated parabolic after bodies of three different thickness ratios. Calcula tion of the flow over the bodies is done by the method of characteristics. The method of Squire and Young is used to calculate the boundary layer properties. Calculation of the pressure distribution on the "mod ified" afterbodies is by Ferri's method of linearized characteristics. N26687* National Gas Turbine Establishment (Gt. Brit.) REFLECTION OF A SMALL PRESSURE PULSE BY DISTRIBUTED FRICTION IN ONEDIMENSIONAL GAS FLOW. P. W. H. Howe. July 1953. 23p. diagrs. (NGTE Memo. M. 168) This work is part of an investigation into the com bustion excited oscillations in a ram jet or reheat system. The main interest here is how much of a pressure pulse produced in the combustion zone succeeds in travelling upstream past the flame stabilizing baffle. The method used is effectively a differentiation from the steady stateconditions and is of general application. NACA RESEARCH ABSTRACTS NO.52 N26694* National Gas Turbine Establishment (Gt. Brit.) " AN IMPROVED DESIGN OF SONIC SUCTION ' PYROMETER. L. Fuller and B. Marlow. June 1953. 24p. diagrs. (NGTE Memo. M. 189) Designs of sonic suction pyrometers for the meas urement of the temperature of low density, high velocity, hot gases have been exmained and tested in a rig that permits the comparison of two pyrometers. one in a gas stream at approximately ground level, atmospheric pressure and the second in the same stream but at a pressure controllable to below 3 inches mercury absolute. Various modifications were made and the pyrometers now advocated agree on mean temperatures to within 50 C when one is at 24 inches mercury absolute and the other is at 3 inches mercury absolute. N26842* Aeronautical Research Council (Gt. Brit.) NOTES ON THE TECHNIQUE EMPLOYED AT THE R. A. E. IN LOWSPEED WINDTUNNEL TESTS IN THE PERIOD 19391945. Edited by F. B. Bradfield. 1952. 71p. photos., diagrs., tab. (ARC R & M 2556; 11,164. Formerly RAE Aero 2222) Very little has been recorded during the war years as to the details of technique used in lowspeed wind tunnel tests. The size and type of tunnel used during this period will remain in use at firms and colleges for some time after newer equipment is available at research establishments, so it has been decided to issue some record of the technique in use at the Royal Aircraft Establihsment during the war years, both with a view to establishing a standard technique where it is satisfactory, and to consider weaknesses where it has failed. N26843* Aeronautical Research Council (Gt. Brit.) THE SOLUTION OF LIFTINGPLANE PROBLEMS BY VORTEXLATTICE THEORY. V. M. Falkner. 1953. 30p. diagrs., 46 tabs. (ARC R & M 2591. Formerly ARC 10,895; S & C 2153; Perf. 354) The report describes in detail the methods by which the principles of vortexlattice theory, introduced in a previous report R & M 1910, are applied to the cal culation of the aerodynamic loading of wings by lift ing plane theory. The scope of the paper is limited to the application of these principles to symmetrical incidence solutions and symmetrical and antisymmel rical wing twist solutions, for which standard solu tions can be treated by comparatively simple loading functions. The effect of discontinuity of direction of leading or trailing edge cannot be avoided even In the simplest solutions, and it has been found necessary to include an investigation of this problem in order to cover the prescribed usage of the method. Special standard functions tabulated in another report are used to allow for the rounding off effects due to change of direction of leading or trailing edge. The general problem of discontinuities is under investi gation and will be dealt with in a later report. A comprehensive set of solutions for a delta wing is in cluded in the report in order to show the convergence NACA RESEARCH ABSTRACTS NO.52 N26843' of and relation between solutions of varying com plexity, and to indicate which solution should be used in order to satisfy the accuracy prescribed for any given problem. The case of the delta wing is not completely general, and the exposition in respect to Induced drag and yawing moment will be completed in a later report. N26847* Aeronautical Research Council (Gt. Brit.) THE EFFECT OF SPANWISE RIBBOOM STIFFNESS ON THE STRESS DISTRIBUTION NEAR A WING CUTOUT. E. H. Mansfield. 1952. 21p. diagrs. (ARC R & M 2663; ARC 11,291. Formerly RAE Structures 13) A theoretical investigation is made into the effect of spanwise ribboom stiffness on the stress distribu tion at a cutout in the interspar skin of a stressed skin wing in bending. Both shear and bending stiff ness of the ribboom are taken into account, and attention is concentrated on the case in which the ribboom is builtin to the spar flanges. Curves are included which determine, for any particular case, the magnitude of the peak shear stress adjacent to the flange, the approximate spanwise variation of this shear stress, the proportion of load transferred by the ribboom to the skin and stringers, and the bend ing moment in the ribboom at its points of attach ment to the spar flanges. By suitable design of the ribboom, it is possible to lower the shear stresses adjacent to the flange with little or no increase in structure weight. N26848* Aeronautical Research Council (Gt. Brit.) WINDTUNNEL TESTS OF THE STALLING PROP ERTIES OF AN 8 PER CENT THICK SYMMETRICAL SECTION WITH NOSE SUCTION THROUGH A POROUS SURFACE. R. C. Pankhurst, W. G. Raymer and A. N. Devereux. 1953. 14p. diagrs., tab. (ARC R & M 2666. Formerly ARC 11,496; Perf. 441; FM 1247) The stalling properties of an 8percentthick sym metrical airfoil with large leadingedge radius of curvature and continuous (distributed) suction over the nose have been tested in the 4foot No. 2 wind tunnel of the National Physical Laboratory. It was found that suction postponed the stall to higher angles of incidence by suppressing separation at the leading edge. The suction also produced beneficial effects in delaying transition. Moreover, it prevented the development of boundarylayer turbulence behind a single excresence or spanwise corrugation, provided the suction was applied over a sufficient chordwise extent of the airfoil surface. The quantity require ments are remarkably small. For example, even at the low Reynolds number of 0.3 x 106 a quantity coefficient CQ(Q Uc) of only 0.0036 is sufficient to increase the lift coefficient at 150 increase by 0.6 (from 0.7 to 1.3), and it is to be expected that CQ will become even less as the Reynolds number is increased. It is not yet possible to estimate the probable power requirements, because the potentiali ties of the best methods of porous construction are not known. 5 N26849* Aeronautical Research Council (Gt. Brit.) THE DETERMINATION OF THE NATURAL FRE QUENCIES OF A FULLSCALE AIRFRAME ENGINE SYSTEM BY THE ADMITTANCE METHOD. J. R. Forshaw and F. T. Mountford. 1953. 24p. diagrs., 3 tabs. (ARC R & M 2667; ARC 11,826. Formerly RAE Tech. Note Structures 20) The development of the method of the measurement of admittances and the solution of the frequency equa tion for a complex fullscale airframeengine system is given, dividing the dynamical system at the attach ment of the engine to the airframe, and using a force system of equal and opposite bending moments and shearing forces. The values of the resonance fre quencies obtained from the graphical solution of the frequency equation and from the resonance test are compared and found to be In good agreement. The method is applicable to the matching of an engine to an airframe by adjusting the flexibility of the mount ing units. N26853* Aeronautical Research Council (Gt. Brit.) HIGHSPEED TUNNEL TESTS OF A 5 PER CENT. CHORD DIVERECOVERY FLAP ON A NACA 0015 AEROFOIL. D. A. Clarke. 1953. 19p. diagrs., 9 tabs. (ARC R & M 2689; ARC 11,743. Formerly RAE Aero 2269) Pressure plotting tests were made in the Royal Air craft Establishment highspeed tunnel on a parallel wooden NACA 0015 wing with diverecovery flap. The Mach number was varied between 0.30 and 0.80, and the Reynolds number was kept constant at 1.4 x 106. All combinations of the following were tested: flap position 0.2c, 0.3c, 0.4c; flap angle 200, 400; Incidence 00, 40. The flapchord/wingchord ratio was 0.05. The report presents a general picture of the action of a diverecovery flap on a wing. The data are, however, too limited to permit the formulation of general design recommendations. N26854* Aeronautical Research Council (Gt. Brit.) THE ROYAL AIRCRAFT ESTABLISHMENT 4 FT x 3 FT EXPERIMENTAL LOW TURBULENCE WIND TUNNEL. PART I GENERAL FLOW CHARACTER ISTICS. H. B. Squire and K. G. Winter. 1953. 28p. diagrs., photos., 4 tabs. (ARC R & M 2690; ARC 10,695; ARC 11,410. Formerly RAE Aero 2182; RAE Tech. Note Aero 1937) The 4 by 3foot wind tunnel was erected as a model of larger tunnels to investigate unconventional design features directed towards obtaining a high standard of flow. Diffusers of 50 cone angle are used, except for the rapid expansion through three wiregauze screens up to the maximum section. The contraction ratio is 31.2:1 and nine screens are fitted in the max imum section. A speed control is used operating in dependently of the fan by means of a bypass duct. The velocity distribution across the workingsection is constant to 1t/4 percent. The standard deviation of the velocity with time measured over a period of 50 see is 0.03 percent. The flow in the diffusers shows no tendency to separate and the velocity dis tribution approaching th# first screen is very satis factory. The installation of cascades with gap 'chord ratio of 1/4 gives uniform outlet flow without apprec  separation in the rapid expansion of the bulge, but the flow in the contraction cone is not satisfactory. A longer contraction would have been advantageous. The power factor has been measured as 0.27 with all screens fitted but could be imporved slightly if all the leaks were sealed. The speed control is satis factory in operation. N26858* Aeronautical Research Council (Gt. Brit.) THE DIFFUSION OF LOAD INTO A PANEL BOUNDED BY CONSTANT STRESS BOOMS AND A TRANSVERSE BEAM. E. H. Mansfield. 1953. 12p. diagrs. (ARC R & M 2729; ARC 11,885. Formerly RAE Structures 31) A theoretical investigation is made into the diffusion of symmetrical, concentrated loads into a long stiffened panel having constant stress edge members and a transverse loading beam. Both pinjointed and clamped end conditions for the beam are considered. Curves are given for determining the peak shear stress near the boom, the variation of this shear stress along the length of the panel, the proportion of load transferred by the beam, and the bending moment at the ends of the beam. N26859* Aeronautical Research Council (Gt. Brit.) AN INVESTIGATION OF THE USE OF AN AUXILIARY SLOT TO REESTABLISH LAMINAR FLOW ON LOWDRAG AEROFOILS. R. W. Cumming, N. Gregory and W. S. Walker. 1953. 14p. diagrs., photos., tab. (ARC R& M 2742. Formerly ARC 13,003; Perf. 645; FM 1424) The use of an auxiliary slot on a laminarflow airfoil has been investigated to check whether laminar flow can be reestablished by suction at the rear of the region of deposited dirt, flies, etc. Results indicate that in the absence of unfavorable pressure gradients, it is possible to reestablish a laminar boundary layer by removing a little more than the whole turbulent layer reaching the slot, and preliminary estimates suggest that with efficient ducting it should be possi ble to achieve a reduction in overall effective drag coefficient by this means. N26861* Aeronautical Research Council (Gt. Brit.) THE EFFECT OF COMPRESSIBLITY ON THE ATTITUDE OF AIRCRAFT IN RECTILINEAR FLIGHT. K. J. Lush. 1953. 7p. diagrs., tab. (ARC R& M 2776; ARC 11,500. Formerly AAEE/ Res/234) The attitude of aircraft (that is, the angle between the aircraft datum and the flight path) is of considerable importance in the aiming of certain airborne arma ment. An investigation was therefore made of the effect of compressibility on the attitude of aircraft in flight in a straight path. The application of the re sults of linear perturbation theory to the problem was examined and the deductions made compared with the results of attitude measurements on a Spitfire IX over a wide range of altitude and airspeed. NACA RESEARCH ABSTRACTS NO.52 N26862* Aeronautical Research Council (Gt. Brit.) THE TOEPLER SCHLIEREN APPARATUS. D. W.\ Holder and R. J. North. 1953. 13p. diagrs., photos., 3 tabs. (ARC R& M 2780. Formerly ARC 13.068; FM 1433) For windtunnel observations it is usually necessary that a Toepler schlieren apparatus shall have roughly uniform sensitivity to density gradients within a cer tain range determined by the nature of the flow. It is also desirable that the illumination in the image of the flow shall be as high as possible if visual obser vations are to be made under the best conditions and if photographs are to be taken with very short expo sures. Methods for satisfying these two require ments are discussed and experimental results are in cluded to illustrate their importance. N26863* Aeronautical Research Council (Gt. Brit.) EVAPORATION OF DROPS OF LIQUID. J. K. Hardy. 1953. 9p. (ARC R & M 2805; ARC 10,675. Formerly RAE Mech. Eng. 1) An analysis has been made of the processes which follow when a drop of liquid is subjected to a sudden change in the condition of the air in which it is suspended. Equations are given from which either the temperature of the drop, or the rate at which it will evaporate, can be calculated. N26886* Royal Aircraft Establishment (Gt. Brit.) TORSIONAL STRENGTH OF BARS IN STEEL, AND ALUMINIUM, MAGNESIUM AND COPPER ALLOYS. E. L. Ripley, A. J. Beard and B. A. J. McCarthy. July 1953. 40p. diagrs., 5 tabs. (RAE Tech. Note Structures 119) Simple empirical relationships between the torsional strength and tensile strength of solid and hollow bars are derived from tests on steel, aluminum alloy, magnesium alloy, and copper alloy bars. Both proof and ultimate conditions of failure are considered. The relationships are used to present design data on the torsional strength of solid and hollow bars. N26887* Royal Aircraft Establishment (Gt. Brit.) SHEAR STRENGTH OF PINS IN STEEL, AND ALUMINIUM, MAGNESIUM AND COPPER ALLOYS. E. L. Ripley and A. J. Beard. July 1953. 28p. diagrs., 5 tabs. (RAE Tech. Note Structures 120) Simple empirical relationships between the ultimate shear strength and ultimate tensile strength of cylindrical pins are derived from tests on steel, aluminum alloy, magnesium alloy, and copper alloy bars. The relationships are used to present design data on the shear strength of pins. NACA RESEARCH ABSTRACTS NO. 52 7 N26888' Royal Aircraft Establishment (Gt. Brit.) LOW SPEED WIND TUNNEL MEASUREMENTS OF THE LIFT ON A 450 SWEPT BACK HALF WING AND CYLINDRICAL BODY. J. A. Lawford. July 1953. 40p. diagrs., 12 tabs. (RAE Tech. Note Aero 2243) Measurements have been made of the lift, drag, and pitching moment on a 450 sweptback half wing mounted on a cylindrical body. Four wing aspect ratios between 0.5 and 1.5 and two body diameters have been tested. For the larger body the lift due to the wing has been analyzed by pressure plotting into contributions due to lift on the wing and to the lift induced by the wing on the body. The measured lift slopes are compared with values calculated by the method of J. Weber (R.A.E. Report No. Aero 2467). The calculation overestimates the lift slope of the wingbody combination by approximately 10 percent. The tests are part of an investigation to establish a method of estimating the lift slope of this configura tion, which occurs when a swept fin is carried on a body of circular section (for example, a jetpipe) and when a single store is carried on a wing tip. N26889* Aeronautical Research Council (Gt. Brit.) INVESTIGATION OF THE INSTABILITY OF A MOVING LIQUID FILM. H. B. Squire. January 25, 1952. 10p. photos., diagrs., tab. (ARC 14,586; CF 213) The stability of a thin layer of liquid moving in still air is studied theoretically with the object of throw ing light on the breakup of films during atomization. It is found that instability occurs if W = T/plU2h< 1 and that the wave length for maximum growth factor, for W 1, is X = 4n7T/p2U2 where p1 is the liquid density, p2 is the air density, U is the film velocity, 2h is the film thickness, and T is the surface tension of the liquid. Comparison with ex perimental data shows fair agreement with the observed wave lengths. N26958* Royal Aircraft Establishment (Gt. Brit.) A TECHNIQUE FOR STUDYING THE BEHAVIOUR OF CINE CAMERA MECHANISMS UNDER DYNAMIC CONDITIONS. G. L. Davies. July 1953. 10p. diagrs., photos. (RAE Tech. Note GW 261) In the design of highspeed cine camera mechanisms, it is not always possible to predict the behavior of the film under dynamic conditions. A simple tech nique has been developed for measuring running speed, register, film acceleration, shutter timing, and other parameters. The camera under test has its normal shutter replaced by a "negative" shutter, and it is then used to photograph a cathode ray oscilloscope which has a linear saw tooth wave form applied to its X deflection plates. The wave shape of the resulting traces on the film can easily be in terpreted to give the required information. NACALangley 111753 4M UNPUBLISHED PAPERS N25367* THE MODERN DEVELOPMENT OF THE THEORY OF PROPELLERS. (Le developpement moderne de la theorie de I'helice). Raymond Siestrunck. September 1953. ii, 137p. diagrs. (Trans. from Institute de Mecanique de la Faculte des Sciences de Paris, 1947) The empirical basis of propeller vortex theory is examined; theories already known are arranged and summarized. The general form of the vortex surfaces with respect to the propeller is specified and evaluation of the field of velocities induced by these vortices is undertaken. The idea is given that the functioning of the propeller depends on the state of the velocities of the surrounding fluid. The problem of determining the distribution of the circu lations which induce the velocities along a blade is studied in detail as are the problems of its adapta tions for practical use. MISCELLANEOUS NACA TN 2135 Errata No. 1 on "THE CALCULATION OF DOWN WASH BEHIND WINGS OF ARBITRARY PLAN FORM AT SUPERSONIC SPEEDS. John C. Martin. July 1950. NACA TN 2590 Errata No. 1 on "CALCULATIONS ON THE FORCES AND MOMENTS FOR AN OSCILLATING WING AILERON COMBINATION IN TWODIMENSIONAL POTENTIAL FLOW AT SONIC SPEED." Herbert C. Nelson and Julian H. Berman. January 1952. UNIVERSITY OF FLORIDA 3 1262 08153 2649 
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