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,'//r National Advisory Committee for Aeroi Research Abstracts OCTOBER 1, CURRENT NACA REPORTS NACA Rept. 997 National Advisory Committee for Aeronautics. SUMMARY OF INFORMATION RELATING TO GUST LOADS ON AIRPLANES. Philip Donely. 1950. ill, 51p. diagrs., photos., 21 tabs. (NACA Rept. 997. Formerly TN 1976) * Available information on gust structure, airplane reactions, and pertinent operating statistics has been examined. This report attempts to coordinate this information with reference to the prediction of gust loads on airplanes. The material covered represents research up to October 1947. NACA Rept. 1000 National Advisory Committee for Aeronautics. CALCULATION OF THE AERODYNAMIC LOADING OF SWEPT AND UNSWEPT FLEXIBLE WINGS OF ARBITRARY STIFFNESS. Franklin W. Diederich. 1950. ii, 29p. diagrs., 10 tabs. (NACA Rept. 1000. Formerly RM LBG27a, TN 1876) A method is presented for calculating the aero dynamic loading, the divergence speed, and certain stability derivatives of swept and unswept wings and tail surfaces of arbitrary stiffness. Provision is made for using either stiffness curves and root rota tion constants or structural influence coefficients in the analysis. Computing forms, tables of numerical constants required in the analysis, and an illustrative example are included to facilitate calculations by means of the method. NACA Rept. 1009 National Advisory Committee for Aeronautics. INVESTIGATION OF FRETTING BY MICROSCOPIC OBSERVATION. Douglas Godfrey. 1951. ii, 10p. photos. (NACA Rept. 1009. Formerly TN 2039) An experimental investigation, using microscopic ob servation and color motion photomicrographs of the action, was conducted to determine the cause of fret ting. Glass and other noncorrosive materials, as well as metals, were used as specimens. A very simple apparatus vibrated convex surfaces in contact with stationary flat surfaces. This led to the con clusion that fretting was initiated by the loosening, due to inherent adhesive forces, of finely divided and apparently virgin material, and that its initiation is independent of vibratory motion or high sliding speeds. The fretting of platinum, glass, quartz, ruby, and mica relegated the role of oxidation as a cause to that of a secondary factor. Fretting occurred to clean nonmetals and metals readily, and glass microscope slides and steel balls provided an excel lent method for visual studies. NACA Rept. 1011 National Advisory Committee for Aeronautics. DYNAMICS OF A TURBOJET ENGINE CON SIDERED AS A QUASISTATIC SYSTEM. Edward W. Otto and Burt L. Taylor, Il. 1951. ii, 12p. diagrs. (NACA Rept. 1011. Formerly TN 2091) A determination of the dynamic characteristics of a typical turbojet engine with a centrifugal compres sor, a sonicflow turbinenozzle diaphragm, and fixedarea exhaust nozzle is presented. A general ized equation for the transient behavior of the engine was developed; this equation was then verified by calculations using compressor and turbine performance charts extrapolated from equilibrium operating data and by experimental data obtained from an engine operated under transients in fuel flow. The results indicate that a linear differential equation for engine acceleration as a function of fuel flow and engine speed for operation near a steady state operating condition can be written. NACA TN 2412 National Advisory Committee for Aeronautics. THEORETICAL FORCE AND MOMENTS DUE TO SIDESLIP OF A NUMBER OF VERTICAL TAIL CONFIGURATIONS AT SUPERSONIC SPEEDS. John C. Martin and Frank S. Malvestuto, Jr. September 1951. 60p. diagrs., photos. (NACA TN 2412) Formulas are obtained by means of the linearized theory for the lateral force due to sideslip, yawing moment due to sideslip, and the rolling moment due to sideslip for norma I tail arrangements consisting of rectangular, triangular, and sweptback vertical tails of arbitrary taper and sweep mounted symmet rically on a horizontal tail of arbitrary shape. The resitlts are restricted to cases where the leading edges are supersonic. The effect of the horizontal tail on the derivatives is evaluated for certain cases. A series of design curves is presented for the de rivatives considered for certain taper ratios. NACA TN 2429 National Advisory Committee for Aeronautics. STUDY OF VORTEX SHEDDING AS RELATED TO SELFEXCITED TORSIONAL OSCILLATIONS OF AN AIRFOIL. Raymond L. Chuan and Richard J. Magnus, California Institute of Technology. September 1951. 49p. diagrs. (NACA TN 2429) Results of the experimental investigation of the self excited torsional oscillation of an NACA 0006 airfoil NO. 8 #AVAILABLE ON LOAN ONLY. ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASH I NGTON 25, D. C., CITING CODE NUMBER, TITLE AND AUTHOR. NACA RESEARCH ABSTRACTS NO.8 suspended elastically are covered. The relationship between torsional oscillation and shedding of vortices was investigated for this airfoil. Two types of oscillation phenomena were found in the investiga tion. One type, exhibited by cases with angles of attack just above stall, persisted with increasing velocity without reaching any apparent limit within the range of velocity attainable in the present wind tunnel. The other type, exhibited by cases with higher angles of attack, only showed selfexcited oscillations in a certain range of velocity, the range decreasing with increasing angle of attack. NACA TN 2439 National Advisory Committee for Aeronautics. A THEORY OF CONDUCTIVITY OF COLDWORKED COPPER. Rolf Landauer. September 1951. 23p. diagrs. (NACA TN 2439) The increase in the resistivity of copper under cold working is calculated. The increase is assumed to be caused by dislocations surrounded by a longrange electrostatic field that scatters the conduction electrons. The amount of scattering is found by the method of deformation potentials of Bardeen and Shockley. The scattering is present in addition to the normal thermal scattering and is regarded as a perturbation in the Boltzmann equation. This per turbation is used to find the incremental resistance per dislocation. From this calculated increment in resistance and the known increase of resistivity of heavily coldworked copper, the number of disloca tions in the coldworked copper is found to be 5 x 1011/square centimeter, in agreement with measurements. NACA TN 2446 National Advisory Committee for Aeronautics. WIDTH OF DEBYESCHERRER LINES FOR FINITE SPECTRAL WIDTH OF PRIMARY BEAM. Hans Ekstein, Armour Research Foundation. September 1951. 9p. diagrs. (NACA TN 2446) Width of a DebyeScherrer line under the combined influence of spectral width of the primary radiation and of the small size of the crystal grains is cal culated, omitting all geometric causes of line broadening. The shape of the crystals is assumed spherical, with the same diameter for all grains. The orientation is assumed to be random, with neg ligible statistical fluctuations. It is shown that the width under these circumstances is the sum of that width which is obtained with very large crystals ("spectral" width) and that which is obtained with monochromatic radiation ("size" width). NACA TN 2448 National Advisory Committee for Aeronautics. XRAY DIFFRACTION BY BENT CRYSTAL LAMELLAE. Hans Ekstein, Armour Research Foundation. September 1951. 14p. diagrs. (NACA TN 2448) A bent crystal lamella may be assumed to consist of individual irregularly shaped blocks having mutual angular disorientation or to be regular in structure except for plastic deformation. The blocks were assumed to be of such a size that the kinematic theory would be valid. Xray integrated intensity was calculated in the first case by adding the inten sities of the individual blocks and in the second case by adding the amplitudes. Theoretical analysis indicated that, in both cases, moderate bending is not expected to change the integrated intensity or to cause broadening of the DebyeScherrer lines. NACA TN 2453 National Advisory Committee for Aeronautics. AN EXPERIMENTAL STUDY OF WATER PRESSURE DISTRIBUTIONS DURING LANDINGS AND PLANING OF A HEAVILY LOADED RECTANGULAR FLATPLATE MODEL. Robert F. Smiley. September 1951. 40p. dlagrs., 3 tabs. (NACA TN 2453) 0 Waterpressure, velocity, draft, wetted length, and acceleration measurements are presented for smoothwater landing and planing tests of a rectan gular flatplate model. Landings were made at fixed trims between 60 and 45, for flightpath angles between 2 and 20, with beamloading coefficients of 18.8 and 36.5. Planing runs were made for trims between 60 and 45.0 For landings, the experimental pressure coefficients based on the equivalent planing velocity appear to be substantially independent of the deceleration of the model. The peak pressures were approximately equal to the dynamic pressure corresponding to the velocity of the peakpressure point. NACA TN 2457 National Advisory Committee for Aeronautics. AIR FORCES AND MOMENTS ON TRIANGULAR AND RELATED WINGS WITH SUBSONIC LEADING EDGES OSCILLATING IN SUPERSONIC POTENTIAL FLOW. Charles E. Watkins. September 1951. 44p. diagrs. (NACA TN 2457) This analysis treats the air forces and moments in supersonic potential flow on oscillating triangular wings and a series of sweptback and arrow wings with subsonic leading edges and supersonic trailing edges. The linearized velocity potential for the wings undergoing sinusoidal torsional oscillations simultaneously with vertical translations is derived in the form of a power series in terms of a frequency parameter. Although as many terms of such a series expansion as may be desired can be deter mined, the terms after the first few become very cumbersome. Closed expressions that include the reduced frequency to the third power, an order which is sufficient for a large class of practical ap plications, are given for the velocity potential and components of chordwise section force and moment coefficients. NACA TN 2462 National Advisory Committee for Aeronautics. INFLUENCE OF REFRACTION ON THE APPLICABILITY OF THE ZEHNDERMACH INTERFEROMETER TO STUDIES OF COOLED BOUNDARY LAYERS. Martin R. Kinsler. September 1951. 39p. diagrs., tab. (NACA TN 2462) NACA RESEARCH ABSTRACTS NO.8 3 An analytical investigation was conducted in order to determine the influence of light refraction on the applicability of the ZehnderMach optical interfer ometer to twodimensional cooledboundarylayer studies. A method is presented for estimating the effects of refraction in a cooled boundary layer. Some numerical results are presented for laminar and turbulent boundary layers over a flat plate having wall to freestream temperature ratios from approxi mately 0.33 to 0.9, a freestream Mach number of 1.0, and a Reynolds number of 5 x 10". NACA TN 2463 National Advisory Committee for Aeronautics. EXPERIMENTAL INVESTIGATION OF THE PRES SURE DISTRIBUTION ABOUT A YAWED CIRCULAR CYLINDER IN THE CRITICAL REYNOLDS NUMBER RANGE. William J. Bursnall and Laurence K. Loftin, Jr. September 1951. 34p. diagrs. (NACA TN 2463) An experimental investigation was made of the pres sure distribution about a circular cylinder at var ious angles of yaw. The results indicate that the flow and force characteristics in the range of Reynolds number based on normal velocity compo nent near and above the critical cannot be determined only by the component of flow normal to the cylinder axis. Localized regions of laminar separation which were present in the 0 yaw case in the supercritical Reynolds number range became less distinct as the yaw angle was increased and completely disappeared for yaw angles of 45 and 600. NACA TN 2464 National Advisory Committee for Aeronautics. TWO AXIALSYMMETRY SOLUTIONS FOR INCOMPRESSIBLE FLOW THROUGH A CENTRIFU GAL COMPRESSOR WITH AND WITHOUT INDUCER VANES. Gaylord 0. Ellis, John D. Stanitz and Leonard J. Sheldrake. September 1951. 34p. diagrs. (NACA TN 2464) Solutions for axially symmetric flow through an impeller with and without inducer vanes were ob tained by relaxation methods using an analysis de veloped in the report. The fluid was considered to be inviscid and incompressible. The impeller select ed was of arbitrary design except that the impeller blades were assumed to have zero thickness and to consist of radial elements. Plots of streamlines, lines of constant velocity, and lines of constant tangential blade force are presented and discussed. NACA TN 2470 National Advisory Committee for Aeronautics. EFFECT OF AN AUTOPILOT SENSITIVE TO YAWING VELOCITY ON THE LATERAL STABILITY OF A TYPICAL HIGHSPEED AIRPLANE. Ordway B. Gates, Jr. and Leonard Sternfield. September 1951. 29p. diagrs., 2 tabs. (NACA TN 2470. Formerly RM L50F22) Calculations have been made to determine the effect of an autopilot sensitive to yawing velocity on the lateral stability of a fighter airplane designed for transonic or supersonic speeds. The effects of inclination of the gyro reference axis and of time lag in the autopilot are discussed. Airplane motions in sideslip subsequent to a disturbance in sideslip are also presented for several representative flight con ditions. NACA TN 2477 National Advisory Committee for Aeronautics. INVESTIGATION OF THE AIRCOMPRESSION PRO CESS DURING DROP TESTS OF AN OLEO PNEUMATIC LANDING GEAR. James H. Walls. September 1951. 17p. diagrs., photo. (NACA TN 2477) A brief study has been made to evaluate the impor tance of the type of aircompression process on the loads produced on an oleopneumatic landing gear during impact and to determine the type of air com pression process actually obtained during drop tests. A simplified analysis to determine the effect which different aircompression processes might have in dicates that the value of the aircompression expo nent should have relatively little effect on the landing gear loads throughout most of the impact. The anal ysis of experimental data obtained in these tests shows that the polytropic exponent ranged from 1.01 to 1.10 for the conditions tested. The general trend of the data appears relatively independent of vertical contact velocity. NACA TN 2478 National Advisory Committee for Aeronautics. A PROCEDURE FOR CALCULATING THE DEVELOPMENT OF TURBULENT BOUNDARY LAYERS UNDER THE INFLUENCE OF ADVERSE PRESSURE GRADIENTS. Kennedy F. Rubert and Jerome Persh. September 1951. 61p. diagrs. (NACA TN 2478) A procedure based on the kineticenergy equation and an extended form of the momentum equation has been devised for calculating the development of turbulent boundary layers in adverse pressure gradients. Predictions, by this method, of turbulent boundarylayer development in comparison with experimental results from several sources are pre sented for a number of cases of flow over flat plates and airfoils and in conical diffusers. In the range of boundarylayer flow short of separation, the agree ment with experiment is, in most cases, quite satis factory; in some instances, however, there experimental results. It is believed, however, that good agreement has been obtained in enough instances to justify continuation of effort along the present lines, particularly with respect to improve ment of the correlations and refinement of the equations. NACA TN 2479 National Advisory Committee for Aeronautics. TABLES OF EXACT LAMINARBOUNDARYLAYER SOLUTIONS WHEN TH E WALL IS POROUS AND FLUID PROPERTIES ARE VARIABLE. W. Byron Brown and Patrick L. Donoughe. September 1951. 68p. diagrs., 2 tabs. (NACA TN 2479) Exact solutions of the laminar boundary equations were computed and tabulated for a range of fixed values of Euler number, temperature ratio, and flow through a porous wall. Euler numbers are 0, 0.5, 1, NACA RESEARCH ABSTRACTS NO.8 and negative values to the separation point. Tem perature ratios are 1, 2, and 4 for the impermeable wall and for two values of coolant flow. In addition, results from temperature ratios of 1/2 and 1/4 are given for the impermeable wall. For each case, boundarylayer thicknesses and heattransfer and friction coefficients were computed and tabulated. NACA TN 2480 National Advisory Committee for Aeronautics. COMPARISON OF HEAT TRANSFER FROM AIRFOIL IN NATURAL AND SIMULATED ICING CONDITIONS. Thomas F. Gelder and James P. Lewis. September 1951. 51p. diagrs., photos., 2 tabs. (NACA TN 2480) An experimental investigation of the heat transfer from an 8footchord airfoil model in clear air and in simulated icing conditions was conducted in the icing tunnel. These results are compared with those obtained in a flight investigation with the same model at similar operating conditions. The tunnel results indicate the effect of tunnel turbulence by the forward movement of transition from laminar to turbulent heat transfer. The flight results indicate that the convective heat transfer in icing is considerably different from that measured in clear air and only slightly different from that obtained in the tunnel during simulated icing. NACA TN 2490 National Advisory Committee for Aeronautics. FLIGHT INVESTIGATION OF SOME FACTORS AF FECTING THE CRITICAL TAIL LOADS ON LARGE AIRPLANES. Harvey H. Brown. September 1951. 119p. diagrs., photos., 2 tabs. (NACA TN 2490) Measurements were made of the control motions and control forces during maneuvers regarded as likely to produce critical tail loads. The pitching motion of the airplane agreed well with that calculated, pro vided the equation of motion was not overly simpli fied. Structural deflections measured during the maneuvers proved to be small. It was found that pilots were inclined to allow the controls to return to neutral at a higher rate than obtained in the initial deflection, producing tail loads of which they had no physical awareness. NACA RM E51G02 National Advisory Committee for Aeronautics. EXPERIMENTAL INVESTIGATION OF FORCED CONVECTION HEATTRANSFER CHARACTERIS TICS OF LEADBISMUTH EUTECTIC. Bernard Lubarsky. September 1951. 30p. diagrs., photo., tab. (NACA RM E51G02) The forcedconvection heattransfer characteristics of leadbismuth eutectic were experimentally inves tigated. Experimental values of Nusselt number for leadbismuth fell considerably below predicted values. The addition of a wetting agent did not change the heat transfer characteristics. NACA RM E51G03 National Advisory Committee for Aeronautics. ESTIMATION OF NEUTRON ENERGY FOR FIRST RESONANCE FROM ABSORPTION CROSS SECTION FOR THERMAL NEUTRONS. Donald Bogart. September 1951. 18p. diagrs., 4 tabs. (NACA RM E51G03) Examination of published data for some 52 isotopes indicates that the neutron energy for which the first resonance occurs is related to the magnitude of the thermal absorption cross section. The empirical relation obtained is in qualitative agreement with the results of a simplified version of the resonance theory of the nucleus of BreitWigner. NACA RM E51G05 National Advisory Committee for Aeronautics. ADAPTATION OF A CASCADE IMPACTOR TO FLIGHT MEASUREMENT OF DROPLET SIZE IN CLOUDS. Joseph Levine and Kenneth S. Kleinknecht. September 1951. 28p. diagrs., photos. (NACA RM E51G05) A cascade impactor, an instrument for obtaining the size distribution of droplets borne in a low velocity air stream, has been adapted for flight cloud droplet studies. Data from two flights are presented. NACA TM 1274 National Advisory Committee for Aeronautics. BEHAVIOR OF FAST MOVING FLOW OF COM PRESSIBLE GAS IN CYLINDRICAL PIPE IN PRESENCE OF COOLING. (K Voprosu o Povedenii Bystrodvizhushchegosya Potoka Szhimae mogo Gaza v Pryamoi Tsilindricheskoi Trube pri Nalichii Okhlazhdenia). G. A. Varshavsky. Sep tember 1951. 8p. diagrs. (NACA TM 1274. Trans. from Zhurnal Tekhnicheskoi Fiziki, v.16, no.4, 1946, p.413416). For compressible flow with friction in a cylindrical pipe the momentum, continuity, and heattransfer equations are examined to determine whether an in crease in Mach number ('thermal' Laval nozzle) is obtainable through heat conduction from the gas through the pipe walls. The analysis is based on the assumption that the wall temperature is negligibly small in comparison with the stagnation temperature of the gas. The analysis leads to a negative result. When the gas cooling is increased by also considering radiation to the wall, a limited region at high temperatures is obtained where Mach number increases were theoretically possible. Obtaining this condition practically is considered impossible. BRITISH REPORTS N10001* Aeronautical Research Council (Gt. Brit.) FURTHER WIND TUNNEL TESTS ON THE EF FECTS OF ICE ACCRETION ON CONTROL CHAR ACTERISTICS. A. Spence. 1951. 12p. diagrs., 2 tabs. (ARC CP 43; ARC 13,259. Formerly RAE Tech. Note Aero 2048) Describes tests on a tailplane with a 9percentthick section, 12 trailingedge angle and aspect ratio of 3 NACA RESEARCH ABSTRACTS NO. 8 to find the effect of reducing the thicknesschord ratio as well as the trailingedge angle on the effects of ice accretion on the control hinge moments. Re sults show that reduction of thickness has little ef fect but that the effects of transition movement on b2(b2 = aCH/a,, where CH is the hingemoment coefficient and i is the flap angle) are of great importance, especially at large trailingedge angles. N10002* Aeronautical Research Council (Gt. Brit.) WIND TUNNEL TESTS ON THE EFFECT OF AC CRETION OF ICE ON CONTROL CHARACTERIS TICS IN TWODIMENSIONAL FLOW. A. S. Halliday, A. S. Batson and D. K. Cox. (Revision of 12,403, S June 9, 1949). May 1, 1950. 18p. diagrs., photo. (ARC CP 42; ARC 13,112) Tests in twodimensional flow were made to provide data for estimating the effect of ice accretion on a tailplane when the trailingedge angle of the airfoil section is changed. The effects of aspect ratio on the results are computed and plotted. Results show that for elevators of 40% chord no advantage can be expected from a finer trailingedge angle; but for a 20% flap the finer trailingedge appears to be supe rior. Larger trailingedge angles, however, give rise to much larger changes in control characteris tics due to normal shifts of transition point on the tailplane and to changes in Reynolds number. N10003* Aeronautical Research Council (Gt. Brit.) A RECORDING SYSTEM FOR FLIGHT TEST DATA. P. A. Hufton, F. G. R. Cook and P. S. Saunders. 1950. 13p. diagrs., photos. (ARC CP 44; ARC 12,925. Formerly AAEE Rept. A.A.E.E..'Res 246) A new recording system for flight test data is pro posed in which the reading of the record is mecha nized. The system employs a transmission system from the instruments which enables the instrument readings to be recorded in the binary digital system. This record is subsequently transcribed in an auto matic "reader" the output of which may be either in the form of tabulated results, or as punched tape suitable for direct use in a sequencecontrolled cal culator. A prototype transmission unit has been constructed. N10004* Aeronautical Research Council (Gt. Brit.) COMPARISON BETWEEN EXPERIMENTAL MEASUREMENTS AND A SUGGESTED FORMULA FOR THE VARIATION OF TURBULENT SKIN FRICTION IN COMPRESSIBLE FLOW. R. J. Monaghan. 1951. 23p. diagrs. (ARC CP 45; ARC 13,260. Formerly RAE Tech. Note Aero 2037) In ARC CP 46, a formula was suggested for the vari ation of turbulent skin friction with Mach number and heat transfer. Experimental results gave a check on the worth of the formula only for flow over a flat plate under zero heattransfer conditions at M = 2.46. The present report extends this check by analyzing data from Univ. of Tex., DRL CM501, covering flow over a flat plate under zero heattransfer conditions for M = 1.9 to M = 2.2, and from NACA RM E8L03, covering subsonic flow through a circular pipe for temperature differences up to 684 C. In each (over) case, a good correlation is obtained on the basis of a known incompressible flow formula. N10005* Aeronautical Research Council (Gt. Brit.) A NOTE ON THE USE OF TIME SERIES IN THE ANALYSIS OF FLIGHT TEST RECORDS. W. P. Jones. 1951. 28p. diagrs. (ARC CP 46; ARC 12,911) Tustin's method for analyzing the behavior of linear systems is briefly reviewed, the numerical analysis being expressed in terms of matrices. Possible ap plications of time series representation to the study of aircraft stability characteristics are discussed, and a detailed numerical investigation of a simple one degree of freedom undamped system is made. For this system the Tustin method of analysis in terms of A units seems satisfactory. An alterna tive method based on the use of Simpson's integra tion rule in conjunction with time series represen tation is also described. N 10006* Aeronautical Research Council (Gt. Brit.) THE EFFICIENCY OF AIRSCREWS ON WINGS OF LARGE CHORD. E. Ower and R. Warden. 1951. 18p. diagrs., 4 tabs. (ARC R & M 2438. Formerly ARC 3945; ARC Ae.Techl. 1509; ARC AP 191) Presents information on the effect on propulsive ef ficiency of varying the distance between the plane of the airscrew and the wing leading edge, when the airscrew diameter is half the wing chord and the wing thickness is sufficient to enclose the engine nacelle. Gives the increase of lift due to the slip stream and compares results with calculations made with the formula of Smelt and Davies. It also gives a theoretical analysis of the effect on propulsive ef fiency of increasing the number of tractor airscrews along the wing leading edge. N10007" Aeronautical Research Council (Gt. Brit.) CALCULATION OF THE INTERFERENCE ON A THIN SYMMETRICAL AEROFOIL WITH HINGED FLAP SPANNING A CLOSED WINDTUNNEL. J. H. Preston and A. R. Manwell. 1951. 14p. diagrs. ARC R & M 2465. Formerly ARC 5388; ARC Ae.1867) Calculations of the tunnel interference corrections for a2 aCLa"!, b1 = aCHi/a', b2 = aCH/an, ml = aCm/aa', m2 = aCm/N (where Cm is the moment coefficient about the leading edge, CH the hinge moment coefficient, q the flap angle, and a the incidence), were made for an infinitely thin symmet rical airfoil with hinged flap, spanning the center of a closed wind tunnel. The method of solution is analogous to that used by Glauert in R & M 1095, and NACA RESEARCH ABSTRACTS NO. 8 includes his results. The corrections for a2, b2, and m2 are roughly equal to or less than that (over) found by previous investigators for the lift, depending on E the flap/chord ratio. The correction for ml is about 50 percent greater and for b1 it is double. N10008* Aeronautical Research Council (Gt. Brit.) PITOTTUBE READINGS NEAR SHOCK WAVES IN THE N.P.L. AND R.A.E. HIGH SPEED TUNNELS. R. G. Fowler. 1951. 13p. diagrs., photos. (ARC R & M 2468. Formerly ARC 8383; ARC FM 773) It is shown that in recent tests in the Royal Aircraft Establishment High Speed Tunnel pitot readings near shock waves agree with theory as well as would be expected. Comparative readings in the National Physical Laboratory 20" x 8" tunnel confirmed that the condensation of atmospheric humidity has a large but unpredictable effect on pitot pressures in this tunnel and in all tunnels of this type. N10009* Aeronautical Research Council (Gt. Brit.) DISTORTION OF CONTROL SURFACE PANELS. D. M. A. Leggett and R. G. Chapman. 1951. 30p. diagrs., photos., 10 tabs. (ARC R & M 2478; ARC 8309; ARC 9991. Formerly RAE SME 3288; SME 3381) Provides data for the design of control surface pan els of any required rigidity, flight experience having shown that distortion of control surface panels, due to air loads, often has a serious effect on the pilot's operating force in the case of large or highspeed aircraft. Part I contains estimates of the distortion that is likely to occur with different types of con struction and for various aerodynamic loading. Part II consists of a detailed account of tests made on various types of control surfaces. N10010* Aeronautical Research Council (Gt. Brit.) WING FLEXUREAILERON FLUTTER TESTS ON A MODEL OF B.A.C. WING TYPE 167. C. Scruton. 1951. 19p. diagrs., photos., 8 tabs. (ARC R & M 2480; ARC 7566; ARC 8106) Test results indicate that symmetrical wing flexure aileron flutter characteristics were slightly more fa vorable with a normal aileron control system than with a proposed flexurally geared aileron system, and the effect of mobility of the fuel in an unbaffled tank was slightly beneficial. It was found also that the minimum damping coefficient necessary to pre vent wing flexureaileron flutter of the B.A.C. wing type 167 varied from a value of 85 lb ft per radian per second at sea level to 420 lb ft per second at an altitude of 30,000 ft. N10011* Aeronautical Research Council (Gt. Brit.) NOTE ON DIFFERENTIAL GEARING AS A MEANS OF AILERON BALANCE. S. B. Gates. 1951. 22p. diagrs. (ARC R & M 2526; ARC 4921. Formerly RAE BA 1642) It has been suggested in some American investiga tions that differential gearing, combined with adjust ment of the aileron floating angle by means of a tab, may be a powerful method of balancing ailerons. This report sets out the theory of this method of bal ance and analyzes it in relation to the most pressing problem of aileron design, which is to obtain close balance at high speed without overbalance in any part of the range, or uncomfortable lightness at slow speed. It is shown that this result can be achieved more directly by differential balance than by any other method if the differential and the tab setting are nicely adjusted to the natural floating properties of the aileron. N10012* Aeronautical Research Council (Gt. Brit.) AN EXPERIMENTAL INVESTIGATION OF THE EF FECT OF LOCALISED MASSES ON THE FLUTTER OF A MODEL WING. N. C. Lambourne and D. Weston. 1951. 28p. diagrs., photos., 2 tabs. (ARC R & M 2533. Formerly ARC 7604) This report contains an account of some experiments on the effect of concentrated masses (representing wing engines, etc.) on the flutter characteristics of a model cantilever wing. Flutter critical speeds and frequencies were measured for an extensive range of mass loading and the results are presented in the form of diagrams. The flutter motions for a few representative conditions of mass loading were de termined by an analysis of cinematograph pictures. The results of experiments on the influence of the flexibility of an engine mounting are also included. N10013* Aeronautical Research Council (Gt. Brit.) CALIBRATION OF THE ROYAL AIRCRAFT ESTAB LISHMENT 24FT WIND TUNNEL. 1. E. Allen and K. V. Diprose. 1951. 17p. diagrs., photo., 6 tabs. (ARC R & M 2566; ARC 5980. Formerly RAE Tech. Note 965) A new calibration of the RAE 24foot wind tunnel is presented. The velocity distribution across the jet in three planes was found at several tunnel speeds. The distribution of static pressure throughout the jet, the relation between the dynamic head at various positions in the tunnel, and the holeinside pressure were investigated. A complete list of all previous calibrations together with results and reasons for discrepancies is included. The plane of reference is taken as 12.5 feet from the jet face, and the mean ve locity over the section is greater than the standard value used up to the present by about 1/2 percent at high speeds and 2 percent at low speeds. N10014* Aeronautical Research Council (Gt. Brit.) A COMPARISON BETWEEN PLAIN AND STRINGER REINFORCED SHEET FROM THE SHEAR LAG NACA RESEARCH ABSTRACTS NO. 8 7 STANDPOINT. M. Fine. 1951. 5p. diagrs. (ARC R &M 2648. Formerly RAE SME 3182) In R. &M's. 2098, 2099, 2100 the stringersheet meth od of solving shear lag problems in stringer rein forced sheet was developed. The present report compares for two simple cases the solution for the plain sheet with that for the stringerreinforced sheet. The solutions are practically identical by the two methods provided that the sheet is considered fully effective in taking end load. This leads to the conclusion that, in regions of tensile stress, at all events, all the skin area is to be included in the stringer area when applying this method. N10120* Royal Aircraft Establishment (Gt. Brit.) A COMPARISON OF LEAD ACID AND ALKALINE BATTERIES FOR USE IN M.T. VEHICLES. J. W. T. Durrell and S. Pomroy. February 1951. 15p. diagrs., tab. (RAE Tech. Note Chem.1144) A comparison has been made of the capacity and voltage characteristics of 24volt M.T. batteries, comprising standard lead acid and alkaline cells ranging from 40 to 110 A.H. capacity at the 10hour rate, at temperatures of +20, 0, and 10 C and at low and high discharge rates. The weight and vol ume of a lead acid battery are always considerably less than those of an alkaline battery of equivalent capacity and high rate voltage performance. The comparison is unchanged even if the batteries are half discharged before being subject to high rate discharges. UNPUBLISHED PAPERS N9423* THE EFFECTIVE WIDTH OF INFINITELY LONG, FLAT RECTANGULAR PLATES UNDER VARIOUS CONDITIONS OF EDGE RESTRAINT. (De meedra gende breedte bij groote overschrijding der knikspanning voor verschillende inklemming der plaatranden). W. T. Koiter. 90p. diagrs., curves, 4 tabs. (Trans. from Nationaal Luchtvaart laboratorium, Amsterdam. S.287, December 31, 1943). This investigation deals with the effective width of (infinitely) long plates with straight, laterally freely movable edges under axial compression. It is as sumed that the proportional limit is not exceeded and that the plate does not buckle between the nails by which it is attached to the stiffeners. The discussion includes hinged, clamped, and elastically restrained edges. In the last case, the torsional rigidity of the stiffeners which produces the edge restraint is taken as independent of the stress and of the wave length of the buckles in the plate. The results ob tained for an infinitely long plate can also be applied to plates of finite but not abnormally small length. N10135* ON THE PROBLEM OF HYDROFOILS FOR SHIP PROPULSION. (Kvoprosu o dvizhenii na podvodnykh kryl'yakh). A. N. Vladimirov. 20p. diagrs., photos. (Trans. from Sudostroenie, v.8, no.7, July 1938, p.411417). In the journal Sudostroenie', No. 2, 1938, there ap peared an article by the engineering designer U. U. Benua on certain problems on the propulsion of ships by hydrofoils. The author gives some historical data, points out the advantages that may be theoreti cally expected from the utilization of hydrofoils, and presents his own proposal along ideas agreeing with the proposal of the French engineer Grunberg. The purpose of the present paper is to extend and deepen somewhat the questions touched upon in the article mentioned and to be of some assistance to those working independently in this field in correctly ori enting themselves and soberly evaluating the actual possibilities of hydrofoils. N 10138" BOUNDED FLOW WITH SEPARATION ABOUT A CIRCULAR CYLINDER. (Otryvnoe Obtekanie Kruglogo Tsilindra v Ogranichennom Potoke). Y. R. Berman. lOp. diagrs. (Trans. from Prikladnaya Matematika i Mekhanika, v.13, no.5, 1949, p.543546). The twodimensional problem of separated flow about a symmetrical contour approximating a circle, the flow being bounded by two parallel walls, is con sidered in this report. The flow scheme and the conditions imposed are the usual ones for the prob lems of the classical flow theory. N10139* NONLINEAR BOUNDARY PROBLEM OF UNSTEADY MOTION OF VISCOUS INCOMPRESSIBLE FLUID. (Nelineinaya Kraevaya Zadacha Neustanovivshegosya Dvizhennia Vyazkoi Neszhimaemoi Zhidkosti). D. E. Dolidze. 27p. (Trans. from Prikladnaya Matematika i Mekhanika, v.12, no.2, Mar.Apr., 1948, p.165180). A viscous incompressible fluid filling an interior or exterior region D bounded by the surface F is as sumed in a state of unsteady motion. The region D is considered simply connected and unchanging in time, and the surface F, regular; that is, surface F is considered to have a continuously varying tangent surface and principal curvature. This report con siders the problem of finding within the region D regular solutions of the NavierStokes equations for given values of the velocity on the boundary and at the initial instant. By constructing the socalled Green Tensor a nonlinear system of inte gral equations is obtained. The solution of this sys tem is effected by the method of successive approxi mations. A sufficient condition of convergence is derived and a uniqueness theorem proved. NACALangtey 10151 2550 UNIVERSITY OF FLORIDA IIIIII I III i11111 IIl II111111 11111 III I 3 1262 09079 7407 .":t 
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