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National Advisory Committee for Aeronautics Research Abstracts NO.33 DECEMBER 3, 1952 CURRENT NACA REPORTS NACA Rept. 1050 FORMULAS FOR THE SUPERSONIC LOADING, LIFT AND DRAG OF FLAT SWEPTBACK WINGS WITH LEADING EDGES BEHIND THE MACH LINES. Doris Cohen. 1951. iii, 40p. diagrs. (NACA Rept. 1050) The method of superposition of linearized conical flows has been applied to the calculation of the aero dynamic properties, in supersonic flight, of thin flat, sweptback wings at an angle of attack. The wings are assumed to have rectilinear plan forms, with tips parallel to the stream, and to taper in the con ventional sense. The investigation covers the moderately supersonic speed range where the Mach  lines from the leadingedge apex lie ahead of the wi. c The trailing edge may lie ahead of or behind theK nh', lines from its apex. The case in which the Nias " from one tip intersects the other tip is not tre cS e Formulas are obtained for the load dlstributial.0 total lift, and the drag due to lift. For the caeii .. which the trailing edge is outside the Mach conefrom ". its apex (supersonic trailing edge), the formulasare complete. For the wing with both leading and tral)ng edges behind their respective Mach lines, a degreeof .. approximation is necessary. It has been found poas  ble to give practical formulas which permit the total'" .. lift and drag to be calculated to within 2 or 3 percent of the accurate linearizedtheory value. The local lift can be determined accurately over most of the wing, but the trailingedgetip region is treated only approximately. Charts of some of the functions derived are included to facilitate computing, and several examples are worked out in outline. NACA Rept. 1063 AIRFOIL PROFILES FOR MINIMUM PRESSURE DRAG AT SUPERSONIC VELOCITIESGENERAL ANALYSIS WITH APPLICATION TO LINEARIZED SUPERSONIC FLOW. Dean R. Chapman. 1952. ii, 14p. diagrs. (NACA Rept. 1063. Formerly TN 2264) A derivation is presented of the basic equations which determine the supersonic airfoil profile having mini mum pressure drag for certain prescribed structural requirements. The basic equations are applicable to a variety of practical structural requirements, and can be used with either linear, second order, or shockexpansion airfoil theory. A solution of the basic equations is found in closed form using linear airfoil theory. The results show that in most cases the optimum profile has a blunt trailing edge. The optimum distribution of thickness depends on the Mach number, airfoilthickness ratio, and base pres sure coefficient. The pressure drag of the optimum profile is compared to that of the biconvex and doublewedge profiles. A graphical method of determining an optimum airfoil is developed and applied to several examples. NACA Rept. 1067 GENERALIZATION OF BOUNDARYLAYER MOMENTUMINTEGRAL EQUATIONS TO THREE DIMENSIONAL FLOWS INCLUDING THOSE OF RO TATING SYSTEM. Artur Mager. 1952. ii, 16p. diagrs. (NACA Rept. 1067. Formerly TN 2310). Boundarylayer equations for application in three dimensional flows are developed. With the use of a fixed velocity profile and an empirical friction law an approximate solution in closed integral form is ob ed for a generalized boundarylayer momentum F t sickness and flow deflection at the wall in the .l lIi case A numerical evaluation of this so lut .7 i. rired out for data obtained in a curving non rouati~g duct shows a fair quantitative agreement ..,vith ire measured values. NACA TN '2809 It EXPERPIENTAL INVESTIGATION OF ECCEN T RICITY RATIO. FRICTION, AND OIL FLOW OF ""SHpQ T JOURNAL BEARINGS. G. B. DuBois and V. W. Ocvirk, Cornell University. November 1952. 79p. diagrs., photos., 4 tabs. (NACA TN 2809) An experimental investigation was conducted to obtain performance data on bearings of lengthdiameter ratios of 1, 1/2, and 1/4 for comparison with theo retical curves. A 1. 375inchdiameter bearing was tested at speeds up to 6000 rpm and with unit loads from 0 to 900 pounds per square inch. Experimental data for eccentricity ratio and friction followed single lines when plotted against a theoretically derived capacity number, which is equal to Sommerfeld number times the square of the lengthdiameter ratio. The form of the capacity number indicates that under certain conditions the eccentricity ratio is theoreti cally independent of bearing diameter. A method of plotting oil flow data as a single line is shown. Methods are also discussed for approximating a maximum bearing temperature and evaluating the effect of deflection or misalinement on the eccen tricity ratio at the ends of the bearings. NACA TN 2813 THEORY AND PROCEDURE FOR DETERMINING LOADS AND MOTIONS IN CHINEIMMERSED HY DRODYNAMIC IMPACTS OF PRISMATIC BODIES. Emanuel Schnitzer. November 1952. 51p. diagrs. (NACA TN 2813) *AVAILABLE ON LOAN ONLY. ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., THE REPORT TITLE AND AUTHOR. t WASHINGTON 25, D. C., CITING CODE NUMBER ABOVE EACH TITLE; r '/ 2 A theoretical method is derived for computing the motions and hydrodynamic loads during water land ings of prismatic bodies involving appreciable im mersion of the chines. A simplified method of com putation covering flatplate and Vbottom bodies with beamloading coefficients greater than unity is given as a separate section. Comparisons of theory with experiment are presented as plots of impact lift coefficient and maximum draftbeam ratio against flightpath angle and as time histories of loads and motions. Fair agreement is found to exist for chine immersed landings for angles of dead rise of 0 and 30, beamloading coefficients from 1 to 36.5, flight path angles from 2 to 90, and trims from 6 to 45. NACA TN 2815 A THEORETICAL INVESTIGATION OF THE EF FECT OF PARTIAL WING LIFT ON HYDHOODYNAM IC LANDINt CHARACTERISTICS OF VBOTTOM SEAPLANES IN STEP IMPACTS. Joseph L. Sims and Emanuel Schnitzer. November 1952. 20p. diagrs. (NACA TN 2815) A theoretical investigation is made of the loads and motions in waterlanding impacts of wide prismatic Vbottom seaplanes for constant partial winglift conditions where the resultant velocity of the sea plane is normal to the keel. An approximate method is given for applying the results of this in'.'esliIgatiin to the more gentr.lt] case of oblique impact. The increase in vertical hdJr.,,J',,njnic load factor due to winglift reduction is shown to be approximately 133 percent of the decrease in air load. NACA TN 2816 WATERPRESSURE DISTRIBUTIONS DUL RINu LANDINGS OF A PRISMATIC MODEL HAVING AN ANGLE OF DEAD RISE OF 221 /2 AND BEAM LOADN'. COEF FICIENTS OF 0.48 AND 0.97. Robert F. Smiley. November 1952. 37p. diagrs. 6 tabs. (NACA TN 2816) As part of an overall program, smoothwater land ing tests of a prismatic float having an angle of dead rise of 221/20 were made. Waterpressure, ve locity, draft, and acceleration data are presented. Landings were made for beamloading coefficients of 0. 48 and 0. 97 at fixed trims between 0. 2 and 30. 3 for a range of flightpath angles from 4. 6 to 25. 9 and also for 90. The experimental pressure distributions are found to be in fair agreement with the predictions of the available theory; however, better agreement is obtained by modification of the theory. NACA TN 2817 A THEORETICAL AND EXPERIMENTAL INVESTI GATION OF THE EFFECTS OF YAW ON PRES SURES, FORCES, AND MOMENTS DURING SEA PLANE LANDINGS AND PLANING. Robert F. Smiley. November 1952. 98p. diagrs., 7 tabs. (NACA TN 2817) A theory for the side force, rolling moment, yawing moment, and pressure distribution during yawed landings and planing of seaplanes was developed. For NACA RESEARCH ABSTRACTS NO.33 the special case of the straightsided wedge without chine immersion, the results of the theoretical analy sis are presented in the form of generalized curves covering all step landing conditions. Experimental impact and planing data are presented for a prismatic wedge having an angle of dead rise of 22.5 and are shown to be in reasonable agreement with the theoret ical predictions. NACA TN 2819 EFFECT OF HIGHLIFT DEVICES ON THE STATIC LATERALSTABILITY DERIVATIVES OF A 45 SWEPTBACK WING OF ASPECT RATIO 4.0 AND TAPER RATIO 0.6 IN COMBINATION WITH A BODY. Jacob H. Lichtenstein and James L. Williams. November 1952. 50p. diagrs., photos., 5 tabs. (NACA TN 2819) This paper contains the results of windtunnel tests to determine the effects of plain and split trailing edge flaps, with and without leadingedge slats, on the lowspeed staticlateralstability derivatives of a 45 sweptbackwingbody configuration. Com parisons between the increments obtained from ex periment and those evaluated by simple sweep theory in combination with measured lift and drag incre ments are also made. NACA TN 2820 AN ANALYSIS OF THE ERRORS IN CURVEFITTING PROBLEMS WITH AN APPLICATION TO THE CAL CULATION OF STABILITY PARAMETERS FROM FLIGHT DATA. Marvin Shinbrot. November 1952. 29p. diagrs., 2 tabs. (NACA TN 2820) The problem of assessing the errors in the param eters obtained from a curvefitting process is con sidered, and a scheme which may be applied toward the solution of such problems is obtained. This method is then specialized to the problem of finding the errors in the calculated stability parameters of an airplane, and an example is given. NACA TN 2821 TORSION TESTS OF ALUMINUMALLOY STIFF ENED CIRCULAR CYLINDERS. J. W. Clark and R. L. Moore, Aluminum Company of America. November 1952. 38p. diagrs., photos., 2 tabs. (NACA TN 2821) Results are presented for the second series of tor sion tests on aluminumalloy stiffened circular cyl inders, the first series having been reported in NACA ARR 4E31. The cylinders were similar in construction except that the wall thickness was 0. 020 inch for the first series and 0. 032 inch for the second series. The significant observation from both series of tests are summarized and some comparisons are made with more recent theoretical work. NACA RESEARCH ABSTRACTS NO.33 NACA TN 2823 LANGLEY FULLSCALETUNNEL INVESTIGATION OF THE MAXIMUMLIFT AND STALLING CHARAC TERISTICS OF A TRAPEZOIDAL WING OF ASPECT RATIO 4 WITH CIRCULARARC AIRFOIL SECTIONS. Roy H. Lange. November 1952. 24p. diagrs., photos. (NACA TN 2823. Formerly RM L7H19) Results are given of an investigation at high Reynolds numbers and low Mach numbers to determine the maximumlift and stalling characteristics of a trape zoidal wing of aspect ratio 4 with 10percentthick, circulararc airfoil sections. The tests included measurements of the lift, the drag, and the pitching moment coefficients of the basic wing and of the wing with 0. 20chord droopnose and trailingedge flaps deflected both alone and in various combinations. Scale effects were investigated at Reynolds number ranging from 3.27 x 106 to 7.67 x 106. NACA TN 2824 EFFECTS OF INDEPENDENT VARIATIONS OF MACH NUMBER AND REYNOLDS NUMBER ON THE MAXIMUM LIFT COEFFICIENTS OF FOUR NACA 6SERIES AIRFOIL SECTIONS. Stanley F. Racisz. November 1952. 32p. diagrs., 2 tabs. (NACA TN 2824) The NACA 65006, 64009, 64210, and 642215 airfoil sections were tested in the Langley low turbulence pressure tunnel to determine the effects of independent variations of Mach number and Reynolds number on the maximumlift character istics for the smooth and rough conditions. The lift characteristics were determined at constant Reynolds numbers ranging from 1. 5 x 106 to 9. 0 x 106 for Mach numbers which varied from 0. 1 to approximately 0. 5 for each Reynolds number. NACA TN 2825 A COMPARATIVE EXAMINATION OF SOME MEAS UREMENTS OF AIRFOIL SECTION LIFT AND DRAG AT SUPERCRITICAL SPEEDS. Gerald E. Nitzberg and Stewart M. Crandall. November 1952. 30p. diagrs. (NACA TN 2825) Systematic trends in the lift and dragcoefficient variation with Mach number for a number of rela tively thick airfoil sections at moderately super critical speeds are pointed out. Shortcomings of transonic similarity rules are discussed and a semi empirical correlation of drag data is presented. Differences between the drag curves for various air foils are reduced by basing drag coefficient on total pressure rather than dynamic pressure. The initial supercritical loss of lift of airfoil sections having supersonic flow over the upper surface only is shown to result primarily from pressure changes on the lower surface. NACA TN 2826 SIMULATION OF LINEARIZED DYNAMICS OF GAS TURBINE ENGINES. J. R. Ketchum and R. T. Craig. November 1952. 25p. diagrs., photo. (NACA TN 2826) 3 A general method of simulation is presented, where by linearized dynamics of gasturbine engines may be simulated in most orderly fashion with greatest economy of equipment and with most direct use of engineering data. Correlation of experimental and simulated responses of a turbojet engine are shown. Use of generalization factors in determining the coefficients necessary for simulation of engine dynamics is discussed. NACA TN 2827 INVESTIGATION OF A DIFFRACTIONGRATINu INTERFEROMETER FOR USE IN AERODYNAMIC RESEARCH. James R. Sterrett and John R. Erwin. November 1952. 36p. photos., diagrs. (NACA TN 2827) A lowcost interferometer that is easy to adjust and has a large field of view is described. This instru ment, which is based on a principle discovered by Kraushaar, uses small diffraction gratings to pro duce and recombine separate beams of light. The usual twoparabolicmirror schlieren system can be converted inexpensively into a diffractiongrating interferometer. Experimental data are presented to verify the ability of the instrument to provide valid and reliable measurements of air density. Photographs of the flow in a supersonic cascade tun nel are included to indicate the quality of the inter ferograms obtained. NACA TN 2828 EFFECT OF A FINITE TRAILINGEDGE THICK NESS ON THE DRAG OF RECTANGULAR AND DELTA WINGS AT SUPERSONIC SPEEDS. E. B. Klunker and Conrad Rennemann, Jr. November 1952. 26p. diagrs. (NACA TN 2828) The effect of a finite trailingedge thickness on the pressure drag of rectangular and delta wings with truncated diamondshaped airfoil sections with a given thickness ratio is studied for supersonic Mach numbers, linearized theory being used to evaluate the surface pressures. In order to facilitate com parison with wings having sharp trailing edges, the position of maximum thickness and base height are determined for least pressure drag as functions of a basepressure parameter. Comparison is then made between the drag of these wings and similar wings with a sharp trailing edge for various aspect ratios and thickness ratios as a iuncuon o01 stream Mach number. The calculations of the drag characteristics for these wings show that significant drag reductions are possible under some conditions at high supersonic speeds. These reductions are relatively independent of aspect ratio for the rec tangular wings but depend considerably on aspect ratio for the delta wings; the smaller aspect ratios show the larger drag reductions. Calculations of the spanwise distribution of drag are included to compare further the effect of a base on the drag for different aspect ratios. 4 NACA TN 2829 EXPERIMENTS ON TRANSONIC FLOW AROUND WEDGES. George P. Wood. November 1952. 34p. diagrs., photos., tab. (NACA TN 2829) Several aspects of transonic flow around the forward portions of wedge profiles were studied by means of interferometry. Measurements were made of the two kinds of flow pattern that occur at the leading edge of a wedge at an angle of attack. The growth of the supersonic region at a sharp convex corner formed by two flat surfaces was observed. The pressuredrag coefficients of a wedge of 14.5 semi angle were measured at Mach numbers of 0.768, 0.819, and 0.854 and were shown to be consistent with those of wedges of smaller angle when plotted according to the transonic similarity law. Conditions at the bases of the shock waves that interacted with boundary layers on the wedge were measured. The method of characteristics was used to calculate the flow behind an experimentally determined sonic line, and the calculated flow field was compared with the measured flow field. The accuracy in the location of the sonic line necessary to give correctly the pres sure distribution on the surface behind it was deter mined. NACA TN 2830 SEVERAL COMBINATION PROBES FOR SURVEYING STATIC AND TOTAL PRESSURE AND FLOW DIREC TION. Wallace M. Schulze, George C. Ashby, Jr. and John R. Erwin. November 1952. 64p. diagrs., photos., tab. (NACA TN 2830) An investigation has been conducted to provide a basis for design of combination probes intended to survey static and total pressure and direction of flow with special reference to subsonic turbomachine testing. Staticpressure probes, yawelement probes, clawtype yaw probes, and combination probes were tested in an 8inchdiameter calibration tunnel at air velocities up to 445 feet per second. From the results of this investigation, the factors which determine the sensitivity of clawtype yaw probes were determined. Satisfactory combination survey probes for sensing static and total pressure and direction of flow in one or two planes were devised. NACA TN 2834 FLOW SURFACES IN ROTATING AXIALFLOW PASSAGES. John D. Stanitz and Gaylord 0. Ellis. November 1952. 31p. diagrs. (NACA TN 2834) In order to investigate the deviation of flow surfaces from their assumed orientation in the usual type of twodimensional solution, threedimensional, in compressible, nonviscous, absolute irrotational fluid motion is determined for flow through rotating axialflow passages bounded by straight blades of finite spacing and infinite axial length lying on meridional planes. Solutions are obtained for five passages with varying blade spacing and hubtip ratio. The results are presented in such a manner as to apply for all ratios of axial velocity to passage tip speed. It is concluded that, for conditions in typical axialflow blade rows, the deviation of flow surfaces NACA RESEARCH ABSTRACTS NO.33 from their assumed orientation in twodimensional solutions is small. NACA TN 2835 EFFECT OF CHANGING PASSAGE CONFIGURATION ON INTERNALFLOW CHARACTERISTICS OF A 48INCH CENTRIFUGAL COMPRESSOR. H  CHANGE IN HUB SHAPE. John Mizisin and Donald J. Michel. November 1952. 35p. diagrs., photo., tab. (NACA TN 2835) The passage contour of a 48inch centrifugal com pressor was modified by changing the shape of the hub to control the deceleration rates along the blade surfaces in order to improve the internal efficiency of the impeller. A comparison of internalflow char acteristics at design flow rate was made with the original impeller and with a modifiedblade impeller that had the same area variation in the passage. In addition, flow characteristics of the modifiedhub impeller over a flow range from maximum flow to near surge at a corrected tip speed of 700 feet per second are presented. At design flow, even though the deceleration rate along the trailing face was less for the modifiedhub impeller than for the original impeller, the clearance losses (increased by the larger ratio of clearance to passage height) caused the totalpressure losses (and thus the relative adiabatic efficiency) to be about the same at the impeller exit for the two configurations, thus pre cluding any improvement in overall performance. As in the original modifiedblade impellers, large losses occurred at the drivingface inlet at negative angles of attack and in regions of large decelerations along the trailingface flow surface. NACA RM 52125 TORSION, COMPRESSION, AND BENDING TESTS OF TUBULAR SECTIONS MACHINED FROM 75ST6 ROLLED ROUND ROD. R. L. Moore and J. W. Clark, Aluminum Company of America. November 1952. 33p. diagrs., photos., 4 tabs. (NACA RM 52125) Tests were made of tubular sections machined from 75ST6 aluminumalloy rolled rod and having ratios of tube diameter to wall thickness D/t ranging from 2 to 150. The purpose of the investigation was to establish curves of strength in torsion, compression, and bending against D/t for the tubular sections and to show to what extent these strengths may be cor related with the mechanical properties of the materi al. In view of the acceptable mechanical properties obtained for the material, the relations obtained be tween these strengths and D/t may be used as a ten tative basis for design of members of the type inves tigated. NACA RM E52109 EFFECT OF RADIANT ENERGY ON VAPORIZA TION AND COMBUSTION OF LIQUID FUELS. A. L. Berlad and R. R. Hibbard. November 1952. 46p. diagrs., 2 tabs. (NACA RM E52109) NACA RESEARCH ABSTRACTS NO.33 The radiative processes involved in combustion were investigated to determine the present role of radiant energy transfer in combustors. It was shown that at present the amount of radiant energy transfer from flame to fuel is quite small in a turbojet combustor. In order to find techniques for making this radiant energy transfer from flame to fuel significant, meth ods of increasing the equivalent graybody emissivi ties of the fuel drops and the flame, as well as the efficiency of the energy transfer itself, were ex amined. The equivalent graybody emissivity of a hydrocarbon fuel may be increased by the use of liq uid or solid, soluble or nonsoluble, additives. It was found that the nonsoluble solid additive was the most desirable one. In addition to the fact that a slurrytype drop may have a greater emissivity, the suspended solid additive may also greatly contribute to the emissivity of the flame. An approximate equa tion was derived from which the equivalent gray body emissivities of such slurrytype drops may be calculated. Data necessary for these calculations were obtained by spectrophotometric analysis of thin slurry films. In addition to the fuel additive, the use of radiationreflecting walls in a combustor was con sidered as a means of maintaining the efficiency of the radiative transfer of energy from flame to fuel drop. It was then shown that more than half the heat of evaporation of the hydrocarbon constituent of this slurrytype drop might be supplied by radiant energy transfer in such a modified turbojet combustor. NACA RM E52126 EXAMINATION OF SMOKE AND CARBON FROM TURBOJETENGINE COMBUSTORS. Thomas P. Clark. November 1952. 12p. photos., 2 tabs. (NACA RM E52I26) Smoke and carbon from turbojetengine combustors were studied by the methods of electron microscopy, chemical analysis, and Xray diffraction. The smoke exhausting from a combustor was found to consist of carbon black, agglomerated into soot. The carbon black had been partially burned in its passage through the flame zone. The smoke result ed from the incomplete combustion of the vaporized fuel; it was not the result of the pyrolysis of fuel droplets. The soft carbon in the dome of the com bustor liner was found to consist of carbon black and soot intermixed with indeterminate complexes such as highboiling fuel ends and partly polymerized and pyrolyzed heavy hydrocarbons. The hard carbon on the walls of the combustor liner was found to be largely a petroleum coke. The coke was apparently formed by the liquid phase cracking, pyrolysis, and subsequent coking on the liner wall of fuel from the spray nozzle. 5 BRITISH REPORTS N17658* Forest Products Research Lab. (Gt. Brit.) THE 'LEVER AID' IN THE KNIFE TEST OF PLY WOOD. P. Burgess. (Forest Products Research Lab.; Reprint from Wood, v. 17, Sept., 1952, p. 340341) This report describes a tool to aid in the conduction of forcible separation or knife tests for assessing the quality of the glue bond between the veneers of plywood. N17728* Royal Aircraft Establishment (Gt. Brit.) MEASUREMENT OF THE MOISTURE CONTENT OF HIGH PRESSURE OXYGEN FOR USE IN AIRCRAFT. II. A HIGH PRESSURE HYGROMETER. M. E. Bedwell and W. G. Glendinning. May 1952. 18p. diagrs., 3 tabs. (RAE Tech. Note Chem. 1175) The gas flow bridge principle has been used in de signing an allmetal hygrometer to measure the moisture content of oxygen gas at transport cylinder pressure. The method could be applied to a produc ing plant, to give a warning if the moisture content of the gas exceeded the maximum permitted in the spec ification. Both glasscapillary and metalcapillary bridges have given satisfactory results at pressures approaching that of the atmosphere. A magnetically operated pressure sensitive device has been made to function at the pressuredifferential available, but would not operate in the bridge because of slight gas leakage past the float, which upset the bridge off balance. N17729* Royal Aircraft Establishment (Gt. Brit.) THE DEVELOPMENT OF FREQUENCY AND VOLT AGE CONTROL SYSTEMS, EMPLOYING TRANS DUCTOR AMPLIFIERS, FOR A 1.75 KVA, 115 V, 1600 C/S MOTORALTERNATOR, WITH SPECIAL REFERENCE TO DYNAMIC CONSIDERATIONS. C. W. Cooper, W. N. Corn and E. W. Eaton. April 1952. 76p. diagrs., photos. (RAE EL. 1480) This report describes the design, development, con struction and performance of Voltage and Frequency control systems for a 1. 75 KVA, 115 V, 1600 c/s MotorAlternator intended for airborne applications. The control amplifiers employ transductors for pow er amplification. Details are given of the static and dynamic performances of the amplifiers, and the transient responses of the control systems. 6 N177 .0 Aeronautical Research Council (Gt. Brit.) FREE STREAMLINE JETS IN SHEAR FLOW, AND THEIR APPLICATION TO THE DESIGN OF SUCTION SLOTS. W. T. Lord. June 20, 1950. 79p. diagrs. (ARC 13, 197; FM 1450) Solutions for the shapes of free streamline jets is suing from a gap between two flat parallel surfaces may serve as a basis for designing slots for sucking away the boundary layer on an airfoil. A brief description is given of some unsuccessful attempts to obtain analytical solutions for such jets when the flat surfaces are collinear and the flow at a large distance from the gap is a constant shearing motion containing vorticity of any in.i'nirude A method of solution for the case when the magnitude of the vorticity is suf ficiently small for terms of the order of its square and higher powers to be neglected is given in detail. This method may be extended to cover the more gen eral case when there is a step between the flat sur faces on each side of the gap. N17751* Aeronautical Research Council (Gt. Brit.) INCOMPIVESSIBLE TWODIMENSIONAL FLOW OF AN INVISC [) FLUID ABOUT A SYMMETRICAL AEROFOIL IN A CHANNEL OR FREE S6 BEAM L. C. Woods. July 7, 1950. 32p. diagrs., 6 tabs. (ARC 13,240; FM 1456) The paper introduces a new and exact method of cal culating the incompressible, inviscid flow about a symmetrical twodimensional airfoil at zero inci dence. The asymmetric case and the effect of cir culation will be dealt with a subsequent paper. The method gives velocities throughout the field and is much quicker than the "relaxation" methods that have been applied to this problem. The principle is to re place the airfoil by a series of small arcs on which it is assumed that the product of the radius of curva ture and the velocity is constant. It is almost equiv alent to replacing the airfoil by a many sided polygon for, at some distance from the airfoil, the field due to a polygon and that due to a profile composed of arcs as defined above, is sensibly the same. Each arc, of course, must cover the same range as the corresponding side of the polygon. The method is termed "The Polygon Method. The idealization takes place in the plane in which the airfoil is rep resented by a slit, that is, the flow plane. The fields due to each of these arcs are added to give the combined field due to the idealized airfoil. The number of arcs selected to replace the airfoil is governed by the accuracy required in the final re sults. Ii the typical example given at the end of the paper, the author found that 12 arcs were sufficient to give accuracies better than 1 percent in the ve locity increment due to the airfoil. In a later paper, the method will be extended to obtain an approximate solution of compressible flow, the accuracy of which is a little better than that of the wellknown KarmanTsien approximation. NACA RESEARCH ABSTRACTS NO.33 417759* Aeronautical Research Council (Gt. Brit.) AN EXACT FOURIER SERIES METHOD OF CALCU LATIN, THE TWODIMENSIONAL INCOMPRESS IBLE FLOW ABOUT AN ASYMMETRICAL AERO FOIL. L. C. Woods. October 12, 1950. 6p. diners. (ARC 13,440; FM 1484; Oxford Univ., LiEnginetring Lab. No. 46) A method of t alc'ulalin the flow about an asymmetric airfoil at incidence in a free stream is developed. It involves the calculation of a number of Fourier series, the constants of which rapidly converge to zero. The flow about an asymmninietric airfoil at in cidence in a channel is tackled by an iterative pro cess which the author expects to be rapidly conver (Irtt Separate solutions are obtained in the two regions separated by the stagnation streamlines and are modified step by step until they match along these streamlines. N178494 National Gas Turbine Establishment (Gt. Brit.) THERMODYNAMIC PROPERTIES OF AIR AND COMBUSTION PRODUCTS OF HYDROCARBON FUELS, PART II. D. Fielding and J. E. C. Topps. July 1952. 72p. diagrs., 7 tabs. (NGTE R. 120) This report is a continuation of NGTE Report No. R 74, civint methods whereby the data required for gas turbine performance calculations, using any fuel or supporter of combustion, may be rapidly computed. The principle of the method given is the addition of a calculated quantity to the corresponding value of the property for dry air, so numerical con sistency is automatically ensured. The theory from which the corrections are deduced is given fully, and functions have been computed permitting determina tion of the true specific heat, the total heat and the entropy function of the combustion products of all common fuelmedium mixtures. The thermody namic aspects of the combustion process are exam ined and parameters for the calculation of combus tion temperature rise are deduced, and calculated functions are given which permit this determination for any fuel and any atmosphere. The data and methods are suitable for the computation of working data sheets, or for the direct determination of gas turbine performance. Examples of the latter are given, also nomographic aids. Special calculations are discussed, and methods are developed whereby data may be prepared for systems involving further elements or compounds, combustible or otherwise, in the fuel or atmosphere. MISCELLANEOUS NACA Rept. 1050 Errata No. 1 on "FORMULAS FOR THE SUPER SONIC LOADING, LIFT AND DRAG OF FLAT SWEPTBACK WINGS WITH LEADING, EDGES BE HIND THE MACH LINES". Doris Cohen. 1951. NACA RESEARCH ABSTRACTS NO.33 DECLASSIFIED NACA REPORTS NACA RM 8107 RESPONSE OF A ROTATING PROPELLER TO AERODYNAMIC EXCITATION. Walter E. Arnoldi, Hamilton Standard Propellers Division, United Air craft Corporation. January 21, 1949. 26p. diagrs. (NACA RM 8107) (Declassified from Restricted, 9/16/52) This report presents a theoretical analysis of the flexural vibration of an aircraft propeller blade sub jected to harmonic aerodynamic exciting forces at a fixed multiple of propeller rotation frequency. The flexural vibration of a rotating propeller blade with clamped shank is analyzed with the object of present ing, in matrix form, equations for the elastic bend ing moments in forced vibration resulting from aero dynamic forces applied at a fixed multiple of rota tional speed. Matrix equations are also derived which define the critical speeds and mode shapes for any excitation order and the relation between critical speed and blade angle. NACA RM E7H21 PRELIMINARY I N'VEST ILuATION OF HYDRAZINE AS A ROCKET FUEL. Paul M. Ordin, Riley 0. Miller and John M. Diehl. May 24, 1948. 35p. diagrs., photos. (NACA RM E7H21) (Declassified from Confidential, 9/16/52) Several properties of hydrazine were investigated. Theoretical calculations were made of performance of hydrazine with liquid 03, liquid 02, H202, and HN03. Hydrazine is poisonous, decomposes some materials, is temperature sensitive, does not propa gate detonation waves, reacts violently with H202, NaMn04, C(N02)4, and HN03 N02, and gives high theoretical performance with several oxidants. NACA RM L6Ll1 WINDTUNNEL INVESTIGATION OF WING INLETS FOR A FOURENGINE AIRPLANE. Walter A. Bartlett, Jr. and Edwin B. Goral. March 11, 1947. 65p. diagrs., photos., 7 tabs. (NACA RM L6Lll) (Declassified from Restricted, 9/16/52) An investigation has been conducted in the Langley propellerresearch tunnel to develop wingleading edge inlets for location between the inboard and out board nacelles on each wing of a fourengine airplane for the Army Air Forces. The investigation includ ed aerodynamic tests of the basic wing and the origi nal inlet, and the development by the NACA of wing inlets for two versions of the airplane. The original inlet was found to decrease the maximum lift coef ficients and to have critical Mach numbers below those of the wing with the basic nose installed. The totalpressure recovery in the oilcooler ducts was poor regardless of the inlet installation. As the sharp expanding bend in this duct cannot be avoided, it is recommended that the oilcooler air be induced through the cowling or from some source other than 7 the subject wing inlet. Two inlets were developed that should be satisfactory for the airplane. The maximum lift coefficients for the model with inlets 5 and 6 installed were about 1. 21 and 1. 22, respect ively, with 0 wing flaps and 1. 87 and 2. 00, respect ively, with 65 wing flaps compared to correspond ing values of 1. 20 and 2. 01 for the model equipped with the faired basic nose. The predicted critical Mach numbers for inlets 5 and 6 for the critical mil itary power highspeed condition for an altitude of 40, 000 feet were 0. 63 and 0. 64, respectively, as compared to 0. 64 for the thickest section of the basic wing. NACA RM L7E20 CORRELATION OF WINDTUNNEL AND FLIGHT DETERMINATIONS OF THE BUFFET SPEED OF AN AIRPLANE EQUIPPED WITH EXTERNAL STORES. H. Norman Silvers and Kenneth P. Spreemann. March 2, 1948. 54p. diagrs., photos., tab. (NACA RM L7E20) (Declassified from Restricted, 9/16/52) Presents a comparison of windtunnel measurements of incremental lift, drag, and pitchingmoment coef ficients due to external stores and measurements of pressures in the vicinity of the external stores with flight determinations of buffet Mach numbers. It is indicated that the variation of the incremental drag coefficient with Mach number of the external store may be used to estimate maximum flight buffet Mach number, and the attainment of local critical pressure in the vicinity of the store is indicative of the lower limit of flight buffet due to stores. NACA RM L7K10 WINDTUNNEL INVESTIGATION OF THE STABIL ITY OF JETTISONED NOSE SECTIONS OF THE D558 AIRPLANE PHASES I AND II. Stanley H. Scher. January 14, 1948. 33p. photos., diagrs., 6 tabs. (NACA RM L7K10) (Declassified from Confidential, 9/16/52) An investigation of the stability of models of the jettisonable nose sections of the D558 airplanes, Phases I and II, has been conducted in the Langley 20foot freespinning tunnel. It was found that both model nose sections tumbled end over end about an approximately horizontal axis. The installation of suitable fins, together with sufficient forward loca tion of the center of gravity, prevented the tumbling and caused the models to descend in a stable nose down attitude. NACA RM L8E11 PRELIMINARY INVESTIGATION TO DETERMINE PROPELLER SECTION CHARACTERISTICS BY MEASURING THE PRESSURE DISTRIBUTION ON AN NACA 10(3)(08)03 PROPELLER UNDER OPERAT ING CONDITIONS. Albert J. Evans and George Liner. July 14, 1948. 52p. Jiarrs., tab. (NACA RM L8Ell) (Declassified from Confidential, 9/16/52) Contains preliminary data obtained by measuring the pressure distribution on the airfoil sections at nine 8 radial stations of an NACA 10(3)(08)03 design pro peller. The results are presented as normalforce and moment coefficients for a range of nominal angle of attack (simple blade element theory) from 0 to 4 for a section Mach number range of about 0. 6 to 1. 15 for the outboard stations and about 0. 3 to 0. 6 for inboard stations. NACA RM L8H11 AERODYNAMIC CHARACTERISTICS OF FLYING BOAT HULLS HAVING LENGTHBEAM RATIOS OF 20 AND 30. John M. Riebe. November 10, 1948. 26p. diagrs., photos., 4 tabs. (NACA RM L8H11I) (Declassified from Restricted, 9/16/52) An investigation of a series of hulls of lengthbeam ratios 6 to 15 previously reported in NACA TN 1305 has been extended to lengthbeam ratios 20 and 30. The hulls of the entire series were designed to have approximately the same hydrodynamic performance with respect to spray and resistance characteristics regardless of lengthbeam ratio. Drag coefficients and longitudinalstability parameters are presented with and without wing interference. Lateralstability parameters are presented with wing interference. NACA RM L9G05 EFFECT OF AN INCREASE IN HULL LENGTH BEAM RATIO FROM 15 TO 20 ON THE HYDRODY NAMIC CHARACTERISTICS OF FLYING BOATS. Arthur W. Carter and Walter E. Whitaker, Jr. August 24, 1949. 29p. diagrs., photos., 2 tabs. (NACA RM L9G05) (Declassified from Restricted, 9/16/52) Investigations of the effect of hull lenthbeam ratio on the hydrodynamic qualities of flying boats, made in Langley tank No. 1, have been extended to include a lengthbeam ratio of 20. The hydrodynamic qual ities determined in this investigation were longitu dinal stability during takeoff and landing, spray characteristics, and takeoff performance in smooth water and takeoff and landing behavior in waves. UNPUBLISHED PAPERS N9269B* Massachusetts Inst. of Tech., Research Lab. for Mechanics of Materials. AN INVESTIGATION OF THE TORSION CREEPTO RUPTURE PROPERTIES OF ALLOY N155. M. E. Shank. (Final rept.) April 4, 1952. 34p. diagrs., photos. (Massachusetts Inst. of Tech., Research Lab. for Mechanics of Materials) This report describes an investigation of the torsion creeptorupture properties of alloy N155. All tests were in the range of 1200 to 1500 F, with times in the 101000 hour range. The effect of stress concentration in the form of transverse cir cular holes was investigated under similar condi tions. The effect of specimen geometry, as affected by slii.ht changes in specimen contours, was found UNIVERSITY OF FLORIDA fll llhl f 1 111111 11111111111111111 111111 llll fll / 11111111 3 1262 09079 7464 NACA RESEARCH ABSTRACTS NO.33 to be great. Preliminary investigation was made in to what appeared to be axial extension of the speci men during torsional creep. N18354* APPLICATION OF TRANSONIC SIMILARITY CON SIDERATIONS TO FLOW ALONG A WAVESHAPED WALL. Carl Kaplan. 1952. 42p. diagrs., tab. (Presented at Eighth International Congress on Theo retical and Applied Mechanics held in Istanbul, Turkey, August 2028, 1952). The PrandtlBusemann smallperturbation method is utilized to obtain the flow of a compressible fluid past an infinitely long waveshaped wall. When the essential assumption for transonic flow (that all Mach numbers in the region of flow are nearly unity) is In troduced, the expression for the velocity potential takes the form of a power series in the transonic similarity parameter. On the basis of this form of the solution, an attempt is made to solve the non linear differential equation for transonic flow past a wavy wall. The analysis utilized exhibits clearly the difficulties inherent in nonlinearflow problems. Nevertheless, by means of the classical method of Integration in series, recursion formulas for certain fourlabelled coefficients are obtained. Remarkably enough, these recursion formulas can be solved analytically and lead to a number of conclusions; for example, that the transonic similarity parameter cannot be greater than fourthirds. In addition, a novel numerical test of convergence, applied to the power series (in the transonic similarity parameter) representing the local Mach number distribution at the boundary, indicates that smooth symmetrical potential flow past the wavy wall is no longer possi ble once the critical value of the undisturbedstream Mach number has been exceeded. NACALangley 12352 4M 
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