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N National Advisory Committee For Aeronautics Research Abstracts 3.59 I ARCH1 9 1954 CURRENT NACA REPORTS NACA Rept. 1123 A STUDY OF INVISCID FLOW ABOUT AIRFOILS AT HIGH SUPERSONIC SPEEDS. A. J. Eggers, Jr., Clarence A. Syvertson and Samuel Kraus. 1953. n. 27p. diagrs., 6 labs. (NACA Rept. 1123. Formerly TN 2646; TN 2729) Steady flow about curved airfoils is investigated analytically, first assuming air behaves as an ideal gas, and then assuming it behaves as a thermally perfect, calorically imperfect gas. Conclusions are drawn from the study. NACA TN 3063 EFFECTS OF WING POSITION AND FUSELAGE SIZE ON THE LOWSPEED STATIC AND ROLLING STABILITY CHARACTERISTICS OF A DELTA WING MODEL. Alex Goodman and David F. Thomas, Jr. February 1954. 66p. diagrs., photos., 3 tabs. (NACA TN 3063) Results are presented of an investigation made to determine the effects of wing position and fuselage size on the lowspeed static and rolling stability characteristics of models having a triangular wing and vertical tail surfaces. Interference increments between the various components are evaluated, and the variation of the verticaltail liltcurve slopes and the efficiency factors with angle of attack as af fected by wing position and body size are presented. TufIgrid pictures of the flow at the vertical tail as affected by wingfuselage interference are also pre sented. NACA TN 3106 AN EVALUATION OF THE SOAPBUBBLE METHOD FOR BURNING VELOCITY MEASUREMENTS USING ETHYLENEOXYGENNITROGEN AND METHANE OXYGENNITROGEN MIXTURES. Dorothy M. Simon and Edgar L. Wong. February 1954. 30p. diagrs., photos., 5 tabs. (NACA TN 3106) A nonaqueous soapbubble method was used to meas ure the burning velocities of some ethyleneoxygen nitrogen and methaneoxygennitrogen mixtures. Burning velocity calculations were based on high speed schlieren motionpicture records of the flame growth and a theoretical expansion ratio. An upper limit in the spatial velocity in the range 2500 to 3500 centimeters per second due to the appearance of rough flames was found for the soapbubble method. Soapbubble burning velocity measurements were compared with measurements by other methods. NACA RM 54A04 .. EFFECTS OF MOLECJLARWWEIGHT ON CRAZING AND TENSILE PROPERTIES OF POLYMETHYL METHACRYLATE. I. Wolock, M. A. Sherman and B. M. Axilrod, National Bureau of Standards. February 1954. lip. diagrs., tab. (NACA RM 54A04) Tensile and crazing properties are reported for five cast polymethylmethacrylate sheets in which the molecular weight of the resin was 90,000, 120,000. 200,000, 490,000. and 3,160,000. respectively. Both stress and stresssolvent crazing tests were conducted. The tensile strength and ultimate elon gation were found to increase rapidly with increasing molecular weight at the lower molecular weights and to level off at molecular weights of approximately 200,000 and 500,000. respectively. There was no change in the modulus of elasticity over the range studied. The specimens with lowest molecular weight broke at low strains without crazing. The,,, stress and strain at which crazing occurred creased with increasing molecular weighWs.j 9 BRITISH REPORTS N28467 / Aeronautical Research Council (Gt. Brit.) KINETIC TEMPERATURE OF WET SURFACES. A METHOD OF CALCULATING THE AMOUNT OF ALCOHOL REQUIRED TO PREVENT ICE, AND THE DERIVATION OF THE PSYCHROMETRIC EQUA TION. J. K. Hardy. 1953. 13p. 2 tabs. (ARC R .i M 2830. Formerly NACA WR A8; NACA ARR 5G131 A method is given for calculating the temperature of a surface wetted either by a pure liquid, such as water, or by a mixture, such as alcohol and vater. The method is applied to the problem of protecting. by alcohol, propellers and the induction system of the engine against ice. The minimum quantity of alcohol required is calculated for a number of arbitrarily chosen conditions. The effect of evapora tion of alcohol is shown by repeating the calculations for a nonvolatile fluid. The method can be applied to other problems in evaporation, for instance, to the evaporation of fuel in the induction system ol the en gine. The psychrometric equation, used in wet bulb hygrometry, is deduced in its general form. The effect of kinetic heating is included in this equation. * 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 U S1ri / f N28468* Aeronautical Research Council (Gt. Brit.) OBSERVATIONS ON A THIN CAMBERED AERO FOIL BEYOND THE CRITICAL MACH NUMBER. E. W. E. Rogers. (Read before VIth International Congress of Applied Mechanics, September 1948) 1953. 16p. diagrs., photos. (ARC R & M 2432. Formerly ARC 13,238; Perf. 685; FM 1455) In the course of surface pressure measurements and wake traverses on an airfoil section of 10 percent thickness chord ratio, tested at high subsonic speeds in the 20 by 8inch (50.8 x 20.3 cm) highspeed wind tunnel of the National Physical Laboratory, it was discovered that at a particular incidence (3.70) an extensive region of supersonic velocity (M = 1.15) existed without the formation of a welldefined shock wave or a rise of drag. The drag coefficient, in fact, decreased markedly as the Mach number was raised from a low value, and this was accompanied by a rearward movement on the upper surface of the posi tion of boundarylayer transition corresponding to a favorable change of the surface pressure gradient. The transition positions measured with the "china clay" method are compared with those estimated from the observed drag coefficients. Direct shadow photographs illustrate the development of the shockwave pattern. N28469* Aeronautical Research Council (Gt. Brit.) A GENERAL TREATMENT OF STATIC LONGITU DINAL STABILITY WITH PROPELLERS, WITH APPLICATION TO SINGLEENGINED AIRCRAFT. E. Priestley. 1953. 20p. diagrs., 2 tabs. (ARC R & M 2732; ARC 7974. Formerly RAE Aero 1944) A general method of treatment of stickfixed static longitudinal stability with propellers is given, distor tion and compressibility effects being neglected. Model fullthrottle data on some singleengined fighters are analyzed for the flapsup condition to establish a basis of estimation of effect of propeller on stability for this type of design. The general effect of propellers on maneuver point, more particularly the effect on Hm Kn, is considered in an appendix. N28470* Aeronautical Research Council (Gt. Brit.) LOWSPEED MEASUREMENTS OF THE PRESSURE DISTRIBUTION AT THE SURFACE OF A SWEPT BACK WING. V. M. Falkner and Doris E. Lehrian. 1953. 36p. diagrs., photo., 27 tabs. (ARC R & M 2741. Formerly ARC 12,647; Perf. 588; S & C 2335) Measurements were made at selected stations on a sweplback wing with and without body. The wing has 450 sweepback, aspect ratio of 3, and constant chord. The section was chosen to be suitable for work at low Reynolds number. This is the first part of a pro gram on a sweptback wing to provide methods of cal culating the pressure at the surface, as well as the general properties of wings. This report covers NACA RESEARCH ABSTRACTS NO. 59 what is thought to be one of the worst types of dis continuity likely to require investigation (excluding deflected flaps); that is, sudden changes in direction of both leading and trailing edges of the wing. ** N28471* Aeronautical Research Council (GI. Brit.) INTERIM REPORT ON VG RECORDS ON HELI COPTERS. H. 1. Birds. 1953. 5p. diagrs., photos. (ARC R ., M 2746; ARC 12,327. Formerly AFEE Rept. Rota 5) Vg records have been obtained during the past year on Hoverfly I helicopters. Some data have also been obtained on a Hoverfly 11 and a Sikorsky S.51. The Vg records on these aircraft were obtained mainly during test flying which included blind flying and some general flying. It was not possible to separate the flight accelerations from the landing accelera tions, but these were small except in the case of engineoff landings which were the subject of separate tests. N28472* Aeronautical Research Council (Gt. Brit.) THE INITIAL BUCKLING OF A LONG AND SLIGHTLY BOWED PANEL UNDER COMBINED SHEAR AND NORMAL PRESSURE. E. H. Brown and H. G. Hopkins. 1953. 19p. diagrs., 4 tabs. (ARC R & M 2766; ARC 12,605. Formerly RAE Structures 42) Recent American experimental work has suggested that the resistance to buckling of wing skin panels under compression or shear loads is improved by aerodynamic suction. A complete theoretical analy sis of this problem is very difficult, because, com pression load necessarily involves the consideration of postbuckling behavior. An approach is made in this report by considering the restricted problem of the initial buckling of a long, thin, and slightly bowed panel under combined shear and normal pressure. The theoretical values of the initial shear buckling stress which agree well with American experimen tal values, increase with both pressure and curva ture; the wave length of the buckles also increases with pressure, but decreases with curvature. The difference between the buckling stresses for simply supported and clamped edges is considerable for a flat panel under shear alone but decreases rapidly with curvature and pressure, thus making the indeterminacy of practical edge conditions of less importance. N28473' Aeronautical Research Council (Gt. Brit.) TESTS ON A SWEPTBACK WING AND BODY IN THE COMPRESSED AIR TUNNEL. C. Salter, C. J. W. Miles and H. M. Lee. 1953. 15p. diagrs., photos., 5 tabs. (ARC R & M 2738. Formerly ARC 13,155; Perf. 671) NACA RESEARCH ABSTRACTS NO. 59 The model was a sweplback wing of symmetrical section and a long cylindrical body. Aspect ratio of 2.39, taper ratio of 0.37, and sweepback of the quarterchord of 42.50 gives the plan form of the wing. The wing was 8.6 percent thick at the root and 10 percent thick at the tip. Results are given of the lift, drag. and pitching moment for angles of attack up to 300. A selection of How pictures is reproduced. N28474* Aeronautical Research Council (Gt. Brit.) TWODIMENSIONAL AEROFOIL DESIGN IN COM PRESSIBLE FLOW. L. C. Woods. 1953. 19p. diagrs., 3 tabs. (ARC R M 2731. Formerly ARC 12,922; FM 1412) This paper deals with the following twodimensional problem. "The design ol an airfoil to give a specified velocity against chord curve at a given freestream Mach number." A "relaxation" method is adopted, based on the differential equations for incompressible and compressible flow. An essential feature of the method is that the calculations are carried out in the (0,*) or wplane in which the air foil is represented by a slit along = 0. The square mesh in this plane is formed by the streamlines (4 = constant), and equipotentials (0 =constant) for incompressible flow about the airfoil. The method is developed for a symmetrical airfoil at zero inci dence, but the modifications necessary for the more general case are indicated. A worked example is given, from which some idea of the accuracy of the method can be gained. The compressible velocity distribution about a known airfoil was taken as the inirutial data. This airfoil was actually 12 percent thick at 30 percent of the chord distance from the leading edge. Using a mesh giving only 14 mesh points on the airfoil, we find that the calculations yield a 12.06 percent airfoil at 28.2 percent of the chord distance from the leading edge. N28475* Aeronautical Research Council (Gt. Brit.) TESTS ON A WHIRLWIND AIRCRAFT IN THE ROYAL AIRCRAFT ESTABLISHMENT 24FT WIND TUNNEL. (Includes: MOMENTUM INVESTIGA TIONS ON FUSELAGEWING INTERFERENCE AND NACELLE DRAG) T. V. Somer'.ille, R. R. Duddy and G. H. L. Buxton. 1953. 17p. diagrs., 8 tabs. (ARC R & M 2603. Formerly RAE BA Dept. Note LWT 30; RAE BA Dept. Note LWT 34) Simple modifications were found to decrease the drag of the airplane. The drag analysis is not complete and is focused chiefly on the drag due to leaks, cooling, and excrescences. A complete record of the tests is given. Modifications which gave an appreci able saving in drag were sealing of leaks and gaps, fairing of exhaust cooling ducts, and fairing of main cooling inlet. The saving in drag corresponds to an increase in maximum speed of about 15 mph. A further saving of 0.8 pound can be obtained by sealing the cartridge chutes. N28476' Aeronautical Research Council (Gt. Brit.) THE INDUCED VELOCITY FIELD OF A ROTOR. K. W. Mangler and H. B. Squire. 1953. 16p. diagrs. (ARC R & M 2642; ARC 11,562; ARC 11,694. Formerly RAE Aero 2247; RAE Tech. Note Aero 1958) A short account and the results of a theoretical in vestigation of the velocity field induced by a lifting rotor are given. The computation is based on the assumptions that the rotor is lightly loaded and that is has an infinite number of blades. This is applied to calculate the induced velocity distribution for disk incidences of 00, 150, 300, 450, and 900. For the downwash at the rotor itself .(the normal component of the induced velocity) the Fourier coefficients are given, as they are needed for the calculation of the blade motion. N28477* Aeronautical Research Council (Gt. Brit.) MEASUREMENTS OF THE AERODYNAMIC DERIVATIVES FOR A HORNBALANCED ELEVATOR. N. C. Lambourne, A. Chinneck and D. B. Betts. 1953. 15p. diagrs., photos., 2 tabs. (ARC R & M 2653. Formerly ARC 12,085; 0.796) This report gives the results of measurements by a forced oscillation method of the direct derivatives (aerodynamic stiffness and damping) for a horn balanced elevator. The tests were made at low air speeds on a complete wingfuselagetail model at 00 and 100 incidence in a wind tunnel. Some informa tion was obtained on the effect of mean elevator angle on the derivatives when the model was at the high incidence. Measurements were also made with trailingedge cords and transition wires in position. The experiments suggest that none of the above fac tors causes a reduction in damping, but the stiffness derivative was found to be considerably influenced by the elevator angle and by the presence of trailing edge cords and transition wires. In general, the measured values are numerically considerably less than those calculated by simple strip theory using twodimensional vortex sheet theory results. N28478* Aeronautical Research Council (Gt. Brit.) MULTIPLEJET WHITESMOKE GENERATORS. C. Salter. 1953. 18p. diagrs., photos. (ARC R & M 2657. Formerly ARC 10,296; FM 1058; ARC 13,004; FM 1425) Descriptions are given of equipment devised for the generation of fairly large quantities of an optically dense white smoke and special attention has been paid to the need for delivering this through long ducts or against an appreciable backpressure. The smoke consists of very small particles of condensed paraffin vapor and is obtained by directing jets of cool air on to highspeed jets of the vapor issuing from very small orifices. The optimum outputs are about 6 cu ft (170 litres) and 8 cu ft (230 litres) per minute from No. 1 and No. 2 generators, respectively, but considerably larger quantities can be delivered 4 with a slight loss of opacity. Under normal operating conditions, the rate of use of paraffin is rather less than 2 cu in. (33 c.c.) (No. 1) and 3 cu in. (49 c.c.) (No. 21 per minute. N28549* Aeronautical Research Council (Gt. Brit.) THE MEASUREMENT OF HEAT TRANSFER AND SKIN FRICTION AT SUPERSONIC SPEEDS. PART IV. TESTS ON A FLAT PLATE AT M = 2.82. R. J. Monaghan and J. R. Cooke. 1953. 42p. diagrs., 3 tabs. (ARC CP 140) This note gives the results of overall heat transfer and boundarylayer measurements made on a flat plate in a 5 inch square supersonic wind tunnel operating at M = 2.82 under atmospheric stagnation pressure conditions. The tests were made to extend the range of results previously obtained at M = 2.43 and used the same experimental equipment. In general, the results confirm those obtained at the lower Mach number and some general conclusions are now drawn concerning the structure and behavior of experimental laminar and turbulent compressible boundary layers on a flat plate. N28550* Aeronautical Research Council (Gt. Brit.) CURVES FOR ESTIMATING THE WAVE DRAG OF SOME BODIES OF REVOLUTION, BASED ON EXACT AND APPROXIMATE THEORIES. L. E. Fraenkel. 1953. 15p. diagrs. (ARC CP 136) Curves are presented for estimating the wave drag, at zero incidence, of forebodies and afterbodies having straight and parabolic profiles. The after bodies are assumed to lie behind an infinitely long cylindrical body. The curves are based on a limited number of exact and secondorder solutions which have been generalized by appealing to the supersonic hypersonic similar ity law and to slender body and quasicylinder solutions. N28551* Aeronautical Research Council (Gt. Brit.) DESIGN OF A RIGHTANGLED BEND WITH CON STANT VELOCITIES AT THE WALLS. A. S. Thom. 1953. 18p. diagrs., 5 tabs. (ARC CP 135) In designing a corner in a twodimensional duct, it is possible, by the insertion of an airfoil, to main tain the same constant velocity on the outer and inner walls. It is, however, necessary to shape these walls to suit the conditions. The present paper gives a method whereby the airfoil and walls can be designed. Two examples are given. N28552* Aeronautical Research Council (Gt. Brit.) AIR COOLING METHODS FOR GAS TURBINE COM BUSTION SYSTEMS. F. J. Bayley. 1953. 44p. diagrs. (ARC CP 133) NACA RESEARCH ABSTRACTS NO. 59 An account is given of the whole of the work which has been done at NGTE on the problem of aircooling gas combustion systems. Each of the different methods of wall cooling is discussed separately and the theory and mechanism of the cooling process is, developed from first principles. It is shown that "sweat, or effusion cooling, is by far the most effective and efficient method, while the use of "louvered* surfaces represents the nearest practical approach to this ideal which is possible while suita ble porous materials remain unavailable. N28553* Aeronautical Research Council (Gi. Brit.) THE MEASUREMENT OF HEAT TRANSFER AND SKIN FRICTION AT SUPERSONIC SPEEDS. PART II. MEASUREMENTS OF OVERALL HEAT TRANS FER AND OF THE ASSOCIATED BOUNDARY LAYERS ON A FLAT PLATE AT MI = 2.43. R. J. Monaghan and J. R. Cooke. 1953. 63p. diagrs.. 3 tabs. (ARC CP 139) A mean value of 0.906 was obtained for temperature recovery factor and the overall heat transfer meas urements from plate to stream agreed well with re sults from the lowspeed formula of reference 2. There was some forward movement of boundary layer transition, a variation in the exponent of the turbulent velocity distribution, and an increase in displacement thickness with heat transfer. However, no decrease in skin friction below its zero heat transfer value was found. N28554* Aeronautical Research Council (Gt. Brit.) THE THEORY OF AEROFOILS WITH HINGED FLAPS IN TWODIMENSIONAL COMPRESSIBLE FLOW. L. C. Woods. 1953. 32p. diagrs. (ARC CP 138) Recently published methods of deducing practical values of the various control characteristics from a knowledge of their theoretical values increases the importance of the theory of twodimensional controls in an inviscid compressible fluid. The classical work of Glauert neglects compressibility and airfoil thickness, and while the more recent work of Goldstein and Preston includes thickness effects, it ignores compressibility. Furthermore, this latter method achieves accuracy for thick airfoils at the cost of a complicated method of calculation. This paper presents a theory of twodimensional controls in compressible flow which is almost as simple to apply as Glauert's theory. An example given by Goldstein and Preston is treated by the author's method to illustrate this point. N28555* Aeronautical Research Council (Gt. Brit.) THE VIBRATIONS OF A SWEPT WING. N. S. Heaps. 1953. 26p. diagrs. (ARC CP 141) The vibrations of a swept wing with ribs parallel to the direction of flight are considered theoretically. The couplings of torsion and flexure due to the skewness of the ribs and the buiidingin of the root section are investigated. NACA RESEARCH ABSTRACTS NO. 59 N28556' Aeronautical Research Council (Gt. Brit.) THE PERFORMANCE OF SOME TYPICAL TURBO JET ENGINE EXHAUST SYSTEMS, WITH PARTICU LAR REFERENCE TO THE EFFECTS OF SWIRL. P. F. Ashwood and P. J. Fletcher. 1953. 28p. diagrs., 3 tabs. (ARC CP 130) Tests made to determine the effect of swirl on the performance of a quarterscale model of a typical turbojet engine exhaust system and propelling nozzle are described. The losses in the system were derived from direct measurement of thrust. The nondimensional thrust, expressed in terms of the nozzle area and the total pressure at inlet to the ex haust diffuser, was found to vary linearly with the ratio of ambient to inlet total pressures. An annular nozzle was found to give slightly over 2 per cent more thrust at the choking condition than the standard system for the same air mass flow. UNPUBLISHED PAPERS N29228' PROPOSED METHOD FOR DEFINING THE TAPER RATIO OF MONOPLANE WINGS. (Una Proposta per la determinazione del rapporto di rastremaxione delle ali monoplane). Giuseppe Gabrielli. February 1954. 14p. diagrs. (Trans. from Onore di Modesto Panetti, 1950, p.6771) A definition is given of taper ratio applicable to all plan forms included within a double infinity of analytically defined wings. A method of plotting the plan form of a wing for whichh area, span, and taper ratio values are given is described. The author de fines an equivalent wing to which static, aerodynamic, aeroelastic, and weight calculations may be referred in first approximation for any wing. DECLASSIFIED REPORTS NACA RM L6L26 FREEFALL MEASUREMENTS AT TRANSONIC VELOCITIES OF THE DRAG OF A WINGBODY CONFIGURATION CONSISTING OF A 450 SWEPT BACK WING MOUNTED FORWARD OF THE MAXI MUM DIAMETER ON A BODY OF FINENESS RATIO 12. Charles W. Mathews and Jim Rogers Thompson. April 2, 1947. 18p. diagrs., photo. (NACA RM L6L26) (Declassified from Confidential, 11 1053) The National Advisory Committee for Aeronautics is measuring drag of a series of complete airplanelike configurations and their various components at tran sonic velocities by the freefall method. This report covers a test of one configuration of this series. The configuration was composed of a 450 sweptback wing of aspect ratio 4.1 mounted forward of the maximum diameter of a 10inchdiameter body of fineness ratio 12 equipped with stabilizing tail fins. The wing has a 5 70inch span and incorporated an NACA 65009 air foil section of 12inch chord perpendicular to the leading edge. The bodytail fin combination was externally identical with a combination tested pre viously by this method. NAGA RM L7C25a TESTS OF A HORIZONTALTAIL MODEL THROUGH THE TRANSONIC SPEED RANGE BY THE NACA WINGFLOW METHOD. Richard E. Adams and Norman S. Silsby. April 11, 1947. 24p. diagrs., photos., tab. (NACA RM L7C25a) (Declassified from Restricted, 11/10/53) A semispan model of the horizontal tail of a fighter airplane was tested at transonic speeds. Measure ments of lift, elevator hinge moment, angle of attack, and elevator angle were made in the Mach number range from 0.75 to 1.04 for elevator de flections ranging from 100 to 10 and for angles of attack of 1.20, 0.40, and 3.40. The hinge moment data are considered to be only qualitative. NACA RM L7D22 SOME PRESSUREDISTRIBUTION MEASUREMENTS ON A SWEPT WING AT TRANSONIC SPEEDS BY THE NACA WINGFLOW METHOD. J. Ford Johnston and Edward C. B. Danforth. June 6, 1947. 21p. diagrs., photos. (NACA RM L7D22) (Declassi fied from Restricted, 11/10/53) First results are given of chordwise pressure distribution measurements on a 450 sweptback wing at transonic speeds. These tests are part of a fundamental investigate ion of flow phenomena near sonic velocity by the NACA wingflow method. Dis tributions were obtained at two spanwise extensions of the halfspan model of 2inch chord and NACA 65210 airfoil section measured perpendicular to the leading edge. The aspect ratios were 2.1 and 3.5. NACALangley 3954 4M UNIVERSITY OF FLORIDA 31 262 j 08 153 0174 
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