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National Advisory Committee for Aeronautics Research Abstracts NO.50 OCTOBER 9, 1953 CURRENT NACA REPORTS NACA Rept. 1073 AN ITERATIVE TRANSFORMATION PROCEDURE FOR NUMERICAL SOLUTION OF FLUTTER AND SIMILAR CHARACTERISTICVALUE PROBLEMS. Myron L. Gossard. 1952. ni, 45p. diagrs.. 9 tabs. (NACA Rept. 1073. Formerly TN 2346) The idea of the iterative transformation procedure suggested by H. Wielandt is explained. Application of the procedure to ordinary naturalvibral ion problems and to flutter problems is shown in numer ical examples. Comparisons of computed results with experimental values and with results obtained by other methods of analysis are, made. NACA Rept. 1077 ..  TWO AND THREEDIMENSIONAI UNSTS BD9 f LIFT PROBLEMS IN HIGHSPEE F Harvard Lomax, Max A. Heaslet, ranklyn B. Fuller and Loma Sluder. 1952. ii, 55p. diagrs., 3 tabs. (NACA Rept. 1077. Formerly TN 2403; TN 2387) The problem of transient lift on two and three dimensional wings flying at high speeds is discussed as a boundaryvalue problem for the classical wave equation. Kirchhoff's formula Is applied so that the analysis is reduced, just as in the steady state, to an investigation of sources and doubles. The appli cations include the evaluation of indicial lift and pitchingmoment curves for twodimensional sinking and pitching wings flying at Mach numbers equal to 0, 0.8, 1.0, 1.2, and 2 0. Results for the sinking case are also given for a Mach number of 0.5. In addition, the indicial functions for supersonic edged triangular wings in both forward and reversed flow are presented and compared with the two dimensional values. NACA Rept. 1080 A THEORETICAL ANALYSIS OF THE EFFECTS OF FUEL MOTION ON AIRPLANE DYNAMICS. AJbertl' t& A. Schy. 1952. ii, 22p. diagrs., 2 tabs. (NakGA Rept. 1080. Formerly TN 2280) The general equations of motion for an airplane with a number of spherical fuel tanks are derived. These equations are applied to two cases with two fuel tanks located in the plane of symmetry. The calculated motions show that the airplane motion may be greatly changed by considering the motion of the fuel and, in particular, that smallamplitude residual oscillations may result. The same type of analysis may be applied to arbitrarily shaped tanks; there *AVAILABLE ON LOAN ONLY. fore, the most general conclusions as to the effects of the fuel motion on airplane dynamics also apply for arbitrarily shaped tanks. NACA Rept. 1090 METHOD FOR CALCULATING LIFT DISTRIBU TIONS FOR UNSWEPT WINGS WITH FLAPS OR AILERONS BY USE OF NONLINEAR SECTION LIFT DATA. James C. Sivells and Gertrude C. Westrick. 1952. ii, 25p. diagrs. 13 tabs. (NACA Rept. 1090. Formerly TN 2283) A method is presented for calculating lift distribu tions for unswept wings with flaps or ailerons uaing nonlinear section lift data. This method is based upon liftingline theory and is an extension to thS method described in NACA.Rep'85... Simpli J fie& computing forms cont4A4ngdlaed.fulinples are ven for both symme(rjirand asymetfacal S J distributions. A fe roknparisons of expIri mental and calculated ah cteristics ar aso. presented. k l3 J 1 NACA Rept. 1101 / FLIGHT INVESTIGATIONOF tE CAL FEEL DEVICE IN AN IRREVERSIBLE ELEATOR CONTROL SYSTEM OF A LARGEMfRPtANE. B. Porter Brown, Robert G. Chilton and James B. Whitten. 1952. ii, 14p. diagrs. (NACA Rept. 1101. Formerly TN 2496) Data are presented showing the flight characteristics of a large airplane having a controlsurface booster and mechanical feel device in the elevatorcontrol system. The tests were made with various force gradients provided by the adjustable feel device. ; The booster was set to operate at a very high boost S, ratio throughout the tests so that the measured or 9 j3pparjnt stickfree stability would be influenced only slightly by the aerodynamic hinge moments. The Sr sultslshow the effect of the feel device on the ha dling qualities of the test airplane and also the design features which should be incorporated in such Mel devices. NACA Rept. 1105 CHORDWISE AND COMPRESSIBILITY CORREC TIONS TO SLENDERWING THEORY. Harvard Lomax and Loma Sluder. 1952. ii, 19p. diagrs., 4 tabs. (NACA Rept. 1105. Formerly TN 2295) Corrections to slenderwing theory are obtained by assuming a spanwise distribution of loading and de termining the chordwise variation which satisfies the appropriate integral equation. Such integral equa tions are set up in terms of the given vertical in duced velocity on the center line or, depending on the ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST, NW, WASHINGTON s25. D. C., CITING CODE NUMBER ABOVE EACH TITLE, THE REPORT TITLE AND AUTHOR. I Of 2 *U.5e 2 type of wing plan form, its average value across the span at a given chord station. The chordwise dis tribution is then obtained by solving these integral equations. Results are shown for flatplate, rectangular, and triangular wings. NACA Rept. 1107 AN EMPIRICALLY DERIVED BASIS FOR CALCU LATING THE AREA, RATE, AND DISTRIBUTION OF WATERDROP IMPINGEMENT ON AIRFOILS. Norman R. Bergrun. 1952. ii, 21p. diagrs., 6 tabs. (NACA Rept. 1107) An empirically derived basis for predicting the area, rate, and distribution of waterdrop impingement on airfoils of arbitrary section is presented. The concepts involved represent an initial step toward the development of a calculation technique which is generally applicable in the design of thermal ice prevent ion equipment for airplane wing and tail surfaces. The calculation technique presented is based on results of extensive waterdrop trajectory computations for five airfoil cases which consisted of 15percentthick airfoils encompassing a moder ate liftcoefficient range. The differential equations pertaining to the paths of the drops were solved by a differential analyzer. NACA Rept. 1109 EXPERIMENTAL INVESTIGATION OF BASE PRES SURE ON BLUNTTRAILINGEDGE WINGS AT SUPERSONIC VELOCITIES. Dean R. Chapman, William R. Wimbrow and Robert H. Kester. 1952. ii, 19p. diagrs., photos., tab. (NACA Rept. 1109. Formerly TN 2611) The pressures acting on the base of blunttrailing edge airfoils have been measured at Mach numbers of 1.25, 1.5, 2.0, and 3.1 and at Reynolds numbers from 0.2 to 3.8 million. Data are presented for 29 profiles both with laminar and with turbulent bound ary layers approaching the trailing edges of the wings. The base pressure is found to be a function primarily of Mach number and the ratio of the bound ary layer thickness at the trailing edge to the trailingedge thickness. NACA Rept. 1111 AN ANALYSIS OF LAMINAR FREECONVECTION FLOW AND HEAT TRANSFER ABOUT A FLAT PLATE PARALLEL TO THE DIRECTION OF THE GENERATING BODY FORCE. Simon Ostrach. APPENDIX B: NUMERICAL SOLUTION OF SIM PLIFIED BOUNDARYVALUE PROBLEM. Lynn U. Albers. 1953. ii, 17p. diagrs., tab. (NACA Rept. 1111. Formerly TN 2635) A formal and general analysis of the freeconvection flow about a flat plate parallel to the direction of the generating body force is made, and velocity and tem perature distributions for Prandtl numbers of 0.01, 0.72, 0.733, 1, 2, 10, 100, 1000, and large Grashof numbers are computed. The distributions for Prandtl number of 0.72 compare favorably with ex perimental values. It is shown that velocities and Nusselt numbers of the same order of magnitude as those associated with forcedconvection flows can be NACA RESEARCH ABSTRACTS NO.50 obtained under freeconvection conditions. A flow and a heattransfer parameter are derived from which the important physical quantities can be com puted. Reasonable agreement is obtained among values of the heattransfer parameter obtained from an approximate theoretical development, experi ments, and the present development. NACA TN 2979 EFFECTS OF SYMMETRIC AND ASYMMETRIC THRUST REVERSAL ON THE AERODYNAMIC CHARACTERISTICS OF A MODEL OF A TWIN ENGINE AIRPLANE. Kenneth W. Goodson and John W. Draper. September 1953. 67p. diagrs., photo., tab. (NACA TN 2979) An investigation was made to determine the magni tude and degree of changes in static forces and moments caused by variations of symmetric and asymmetric thrust reversal on a twinengine air plane. The effects of both positive and negative thrust coefficients were investigated on the Indi vidual propellers and covered a thrustcoefficient range of 0.167 to 0.150. NACA TN 2983 LINEARIZED POTENTIAL THEORY OF PRO PELLER INDUCTION IN A COMPRESSIBLE FLOW. Robert E. Davidson. September 1953. 47p. diagrs., 5 tabs. (NACA TN 2983) This paper gives the potentialtheory representation of the waveequation flow about a liftingline pro peller of finite number of blades and arbitrary cir culation distribution. From the velocity potential, the compressible inflow velocities at the blade be came known. The induced velocities are known also at any point in the flow because the velocity potential is determined for the whole field. NACA TN 2997 APPLICATION OF SEVERAL METHODS FOR DE TERMINING TRANSFER FUNCTIONS AND FRE QUENCY RESPONSE OF AIRCRAFT FROM FLIGHT DATA. John M. Eggleston and Charles W. Mathews. September 1953. 74p. diagrs., 2 tabs. (NACA TN 2997) A study is presented of several methods for deter mining the transfer functions and frequency response of aircraft from flight tests. Results obtained from experience in the use of these methods are com pared as to time required for application, compara tive accuracy, and means for facilitating their use. The studies cover three categories of methods: sinusoidal response, Fourier analysis of transients, and curvefitting analysis of transients Three general types of aircraft are used to illustrate the application of these methods. NACA TN 2999 IMPINGEMENT OF DROPLETS IN 900 ELBOWS WITH POTENTIAL FLOW. Paul T. Hacker. Rinaldo J. Brun and Bemrose Boyd. September 1953. 58p. diagrs., 2 tabs. (NACA TN 29991 NACA RESEARCH ABSTRACTS NO.50 Trajectories were determined for droplets in air finiii, t&hroupi 900 elbows especially designed for twodimensional potential motion with low pressure losses. The elbows were established by selecting as walls of each elbow two streamlines of the flow field produced by a complex potential function that establishes a twodimensional flow around a 900 bend. An unlimited number of elbows with slightly differ r ti shapes can be established by selecting different pairs of streamlines as walls. The elbows produced by the complex potential function selected are suitable for use in aircraft airintake ducts. The droplet impingement data derived from the trajectories are presented along with equations in such a manner that the collection efficiency, the area, the rate, and the distribution of droplet im pngernement can be determined for any elbow de fined by any pair of streamlines within a portion of the flow field established by the complex potential function. Coordinates for some typical stream lines of the flow field and elocirt, components for several points along these streamlines are pre sented in tabular form. NACA TN 3005 HEAT TRANSFER AND SKIN FRICTION BY AN INTEGRAL METHOD IN THE COMPRESSIBLE LAMINAR BOUNDARY LAYER WITH A STREAM WISE PRESSURE GRADIENT. Ivan E. Beckwith. September 1953. 55p. diagrs., tab. (NACA TN 30051i A simplified method has been developed for the cal culation of heat transfer and skin friction in the compressible laminar boundary layer with an arbi trary Prandtl number near unity and an arbitrary streamwise pressure gradient and wall temperature distribution. The use of a fifthdegree polynomial for the stagnation enthalpy profile gives results of good accuracy for the case of boundarylayer cooling. The method has been extended to the calcu lation of heat transfer under conditions of equili brium dissociation in the boundary layer. By means of a suitable transformation the method is also applied to a body of revolution with a boundary layer thickness of the order of the body radius. NACA TN 3007 LIFT AND PITCHING MOMENT AT LOW SPEEDS OF THE NACA 64A010 AIRFOIL SECTION EQUIPPED WITH VARIOUS COMBINATIONS OF A LEADINGEDGE SLAT, LEADINGEDGE FLAP, SPLIT FLAP, AND DOUBLESLOTTED FLAP. John A. Kelly and NoraLee F. Hayter. September 1953. 45p. diagrs., photos., 2 tabs. (NACA TN 3007) Results of a twodimensional windtunnel investiga tion of an NACA 64A010 airfoil equipped with a leadingedge flap, a leadingedge slat, a split flap, and a doubleslotted flap are presented. The re sults include determination of opt lnunm slat posi tions and effects of varying Reynolds number on the lift and pitchingmoment characteristics of the model with the various highlift devices. 3 NACA TN 3008 EFFECTS OF FINITE SPAN ON THE SECTION CHARACTERISTICS OF TWO 450 SWEPTBACK WINGS OF ASPECT RATIO 6. Lynn W. Hunton. September 1953. 32p. diagrs. (NACA TN 3008. Formerly RM A52A10) A study has been made of the Finite s.pri effects on the local loading characteristics of two siepitrack wings at low speed with a view toward providing some insight into the usefulness of twodimensional section data and spanloading theory for determining the sec tion characteristics of a swept wing. The two wings considered were identical in plan form having 450 of sweepback of the quarterchord line, an aspect ratio of 6, and a taper ratio of 0.5 but differed in twist and in sections, the latter being the NACA 64A010 and NACA 64A810. The analysis is based on compari sons of local pressure distributions and local lift characteristics on the wings with comparable two dimensional section data, all of which were available at large scale. NACA TN 3009 VELOCITY POTENTIAL AND AIR FORCES ASSOCIATED WITH A TRIANGULAR WING IN SUPERSONIC FLOW, WITH SUBSONIC LEADING EDGES, AND DEFORMING HARMONICALLY ACCORDING TO A GENERAL QUADRATIC EQUA TION. Charles E. Watkins and Julian H. Berman. September 1953. 61p. diagrs., tab. (NACA TN 3009) The velocity potential for a triangular wing with sub sonic leading edges experiencing harmonic deforma tions in supersonic flow is treated herein. The oscillations considered are such that the amplitude of distortion of the wing can be represented by a general quadratic equation. The velocity potential is derived in the form of a power series in terms of the fre quency of oscillation. Although only the first four terms of the series expansion are presented, addi tional terms may be obtained if desired. The mate rial constitutes an extension of the work given in NACA Report 1099. NACA TN 3011 COEFFICIENT OF FRICTION AND DAMAGE TO CONTACT AREA DURING THE EARLY STAGES OF FRETTING. I GLASS, COPPER, OR STEEL AGAINST COPPER. Douglas Godfrey and John M. Bailey. September 1953. 23p. diagrs., photos., 2 tabs. (NACA TN 3011) Experiments were conducted to measure the coeffi cient of friction 1i and to determine the damage to the contact area during early stages of fretting of copper at a frequency of 5 cycles per minute. Specimen combinations of copper against glass, copper against copper, and copper against steel, as well as various copper oxide films and powder com pacts, were used. The results lead to the conclu sion that fretting of copper starts with the same mechanical damage that occurs during unidirectional sliding. Fretting of copper against glass, copper against copper, and copper against steel starts with adhesion and metal transfer (galling) with accompa nying high g values (>1.0) the same as those ob tained during unidirectional sliding. After the ini 4 tial high values of ji, a reduction in 1 was ob served, associated with reduced plowing and an in creasing concentration of debris in and around the contact area. After approximately 100 cycles of fretting, gi reached a constant value (0.50.6) ap proximately the same as that obtained with com pacts of either cuprous or cupric oxide. The presence of preformed cuprous or cupric oiide films on copper does not delay the occurrence of fretting but only lowers the initial coefficient of friction. NACA TN 3012 AN ANALYSIS OF TURBOJETENGINEINLET MATCHING. DeMarquis D. Wyatt. September 1953. 19p. diagrs. (NACA TN 3012) A method of presenting turbojetengine airflow requirements and inletsystem airflow capacities in identical though independent parametric terms is developed. The application of the airflow repre sentation technique to the analysis of engineinlet matching conditions is demonstrated. Several ex amples are presented to illustrate the application of the method to the explicit determination of inlet geometric variations required to improve the power plant performance of supersonic airplanes. NACA TN 3014 CALCULATED SPANWISE LIFT DISTRIBUTIONS AND AERODYNAMIC INFLUENCE COEFFICIENTS FOR UNSWEPT WINGS IN SUBSONIC FLOW. Franklin W. Diederich and Martin Ziotnick. September 1953. 120p. diagrs., 11 tabs. (Tables of F matrices to be used with TN 3014 are published separately) (NACA TN 3014) Spanwise lift distributions have been calculated for nineteen unswept wings with various aspect ratios and taper ratios and with a variety of angleof attack or twist distributions, including flap and aile ron deflections, by means of the Weissinger method with eight control points on the semispan. Also calculated were aerodynamic influence coefficients which pertain to a certain definite set of stations a long the span, and several methods are presented for calculating aerodynamic influence coefficients for stations other than those stipulated. NACA TN 3015 AN EXPERIMENTAL INVESTIGATION OF SECOND ARY FLOW IN AN ACCELERATING, RECTANGULAR ELBOW WITH 900 OF TURNING. John D. Stanitz, Walter M. Osborn and John Mizisin. October 1953. 60p. diagrs., photos., 2 tabs. (NACA TN 3015) Secondary flow tests were conducted on an accelerat ing elbow with 900 of turning designed for prescribed velocities that eliminate boundarylayer separation by avoiding local decelerations along the walls. Sec ondary flows were investigated for six boundarylayer thicknesses generated on the plane walls of the elbow by spoilers upstream of the elbow inlet. The passage vortex associated with secondary flows appears to be near the suction surface and away from the plane wall of the elbow at the exit and does not have appre ciable spanwise motion as it moves downstream from the elbow exit. As the spoiler size increases, the boundarylayer form changes and a rather sudden NACA RESEARCH ABSTRACTS NO.0S  difference in the secondary flow occurs, perhaps associated with the reduced importance of viscous effects in thick boundary layers. It is suggested that the strength of the secondary vortices is small and that the energy of secondary flows is small. NACA TN 3016 ANALYSIS OF TURBULENT HEAT TRANSFER AND FLOW IN THE ENTRANCE REGIONS OF SMOOTH PASSAGES. Robert G. Deissler. October 1953. 88p. diagrs. (NACA TN 3016) A previous analysis for fully developed turbulent heat transfer and flow with variable fluid properties is extended and applied to the entrance regions of smooth tubes and parallel flat plates. Integral hbeat transfer and momentum equations are used for calcu lating the thicknesses of the thermal and flow bound ary layers. The effect of variable properties is determined for the case of uniform heat flux, uniform initial temperature distribution, and fully developed velocity distribution. A number of other cases In which the fluid properties are,constant are analyzed. The predicted Nusselt numbers for air with a uniform wall temperature and uniform initial temperature and velocity distributions agree closely with experimen tally determined values. NACA TN 3017 AXIALLOAD FATIGUE TESTS ON NOTCHED AND UNNOTCHED SHEET SPECIMENS OF 61ST6 ALUMINUM ALLOY, ANNEALED 347 STAINLESS STEEL, AND HEATTREATED 403 STAINLESS STEEL. Herbert F. Hardrath, Charles B. Landers and Elmer C. Utley, Jr. October 1953. 28p. diagrs., 4 tabs. (NACA TN 3017) Axialload fatigue tests at a stress ratio of zero were performed on notched and unnotched sheet specimens of 61ST6 aluminum alloy and 347 and 403 stainless steels. Special emphasis was placed on tests at high stress levels producing failures in small numbers of cycles. It was found that the stressconcentration factors effective in fatigue of notched specimens were somewhat less than the theoretical elastic values at low stresses and were approximately equal to one at the ultimate strength. The mimimum life to failure at stresses near the ultimate strength was drastically reduced with m creasing stressconcentration factor. NACA TN 3019 INVESTIGATION OF THE STATISTICAL NATURE OF THE FATIGUE OF METALS. G. E. Dieter and R. F. Mehl. Carnegie Institute of Technology. September 1953. 25p. diagrs., 5 tabs. (NACA TN 3019) Results are presented of an investigation of the sta tistical nature of the fatigue of metals utilizing sta tistical methods developed previously. The inves tigation included a study of the fatigue properties and their statistical variation of a plain carbon eutectoid steel heattreated to coarse pearlitic and spheroi dized structures of the same tensile strength and of commercially pure aluminum (2S) and 245 alloy heattreated to two different structures. Calcula NACA RESEARCH ABSTRACTS NO.50 tion of data from the literature provided statistics for 75S aluminum alloy for comparison with the data of the present investigation. NACA RM E53E05 EFFECT OF PRESSURE ON THE SMOKING TEND ENCY OF DIFFUSION FLAMES. Rose L. Schalla and Glen E. McDonald. September 1953. 13p. diagrs., 2 tabs. (NACA RM E53E05) The effect of pressure on smoke formation was in vestigated by burning nine hydrocarbon fuels as diffusion flames from a modified wick lamp in an en closed chamber. The maximum relative rate at which each fuel could be burned without smoking was determined over a pressure range of about 1/2 to 4 atmospheres, and up to 12 atmospheres in the case of one fuel. The results indicate that over this pressure range the maximum smokefree fuel flow was inversely proportional to the pressure. The relative variation in smoking tendency among the different fuel types was approximately constant at all pressures. From an analysis of the data it has been tentatively proposed that the variations in smoke formation with pressure result from changes in the rate of diffusion and mixing of the fuel and air. NACA RM E53G08 VAPOR PRESSURES OF CONCENTRATED NITRIC ACID SOLUTIONS IN THE COMPOSITION RANGE 83 TO 97 PERCENT NITRIC ACID, 0 TO 6 PER CENT NITROGEN DIOXIDE, 0 TO 15 PERCENT WATER, AND IN THE TEMPERATURE RANGE 200 TO 800 C. A. B. McKeown and Frank E. Belies. September 1953. 22p. diagrs., tab. (NACA RM E53G08) Total vapor pressures were measured for 28 acid mixtures of the ternary system nitric acid, nitrogen dioxide, and water within the temperature range 200 to 800 C and within the composition range 83 to 97 percent nitric acid, 0 to 6 percent nitrogen dioxide, and 0 to 15 percent water. The ullage of the appara tus used for the measurements was 0. 65. Ternary diagrams showing isobars as a function of composi tion of the system NO2H20HNO3 were constructed from experimental and interpolated data for the temperatures 250, 400, and 600 C and are presented herein. NACA RM L53G24a MODEL DITCHING INVESTIGATIONS OF THREE AIRPLANES EQUIPPED WITH HYDROSKIS. (Revised) Lloyd J. Fisher. September 1953. 8p. photos. (NACA RM L53G24a) Calmwater tests were made to determine possible arrangements of hydroski ditching gear on typical multiengine airplanes. The tests showed that hydro skis would afford very satisfactory water landings as compared with landings without skis. The best Landings were made in a nearlevel (slightly nose up) landing attitude although any normal landing attitude was satisfactory. It is possible that critical damage could be eliminated from ditchings by using a hydroski landing gear. 5 BRITISH REPORTS N11456* Aeronautical Research Council (Gt. Brit.) THE DISTURBED MOTION OF ARTICULATED BLADES. H. Roberts. October 1949. 77p. diagrs. (ARC 12, 688; H. 127) The general theory of motion for articulated heli copter blades is presented and allowance is made for for: (a) arbitrary motion of the hub, (b) offset blade hinges, (c) inclined flapping hinges, and (d) cam bered blades. The method employed is the simple one of giving the hub prescribed horizontal, vertical, and angular velocities, the motion being two dimensional. It is shown that the flapping stability equation arises when considering disturbed motion or response phenomena. It is also shown that dis turbances in the pitching plane lead in general to a lateral tilt of the rotor disk, thereby inducing a coupling between the longitudinal and lateral stability equations. N26520 Aeronautical Research Council (Gt. Brit.) THE DETERMINATION OF SKIN TEMPERATURES ATTAINED IN HIGH SPEED FLIGHT. F. V. Davies and R. J. Monaghan. 1953. 65p. diagrs., 8 tabs. (ARC CP 123) This report discussed the factors affecting skin temperatures attained by highspeed missiles and presents some methods of solution. These have been reduced to graphical or tabular form and are set out in order of complexity. Graphical or alge braic solutions may be quickly obtained if steady con ditions are assumed, and for some flight cases these are reasonable approximations to correspond ing transient solutions. If the temperature time variation is required, then longer numerical integra tion processes have to be performed. Account may be taken of external radiation and heat loss to the in terior if their effects are considered significant. N26521* Aeronautical Research Council (Gt. Brit.) THE DEFINITIONS OF THE ANGLES OF INCI DENCE AND OF SIDESLIP. C. H. E. Warren. 1953. lip. diagrs. (ARC CP 124) The use of large angles of incidence and of sideslip in missile work, and recent changes in windtunnel testing techniques, .have shown the need for clear and precise definitions of the angles of incidence and of sideslip. The suitability of different definitions for both experimental and theoretical work in both the aircraft and missile field is considered, and it is concluded that, as no single definition is univer sally acceptable, care should be taken in theoretical and experimental reports to define precisely the angles used. 6 N26522" Aeronautical Research Council (Gt. Brit.) PRESSURE ERROR MEASUREMENT USING THE FORMATION METHOD. K. C. Levon. 1953. 15p. diagrs. (ARC CP 126) Measurements of pressure error at altitude have been made by flying several aircraft in formation with a reference aircraft whose airspeed system had previously been calibrated by radar. Tests made show that analysis by comparison of indicated air speeds (comparison of differences between static and total head pressures) gives more consistent and reliable results than pressure altitude comparison (comparison of direct measurement of static pres sure). N 26523* Aeronautical Research. Council (Gt. Brit.) SURFACE SLOPES AND CURVATURES OF THE RAE 100 104 AND OTHER ROOFTOPTYPE AEROFOIL SECTIONS. J. Williams and Edna M. Love. 1952. lip. 3 tabs. (ARC CP 129) Formulas and tables have been obtained for the accurate and rapid calculation of the surface slopes and curvatures of the RAE 100104 airfoil sections, and of more general "rooftoptype" sections. In particular, the surface slopes of the RAE 102 and 104 shapes have been evaluated for application with a "tangentplane" method of model construction. N26535* Aeronautical Research Council (Gt. Brit.) THE EFFECT OF COMPRESSIBILITY ON THE PERFORMANCE OF A GRIFFITH AEROFOIL. H. H. Pearcey and E. W. E. Rogers. 1953. 30p. diagrs., photos., 4 tabs. (ARC R & M 2511. Formerly ARC 10, 096; FM 1017; Perf. 250) Experiments were made in the 20inch by 8inch highspeed tunnel at the NPL on a 9inch chord, 22 percent thick, symmetrical Griffith section at 00 incidence. Drag was determined by the pitot traverse method. Information on the flow was ob tained from the pitot tube traverses, from direct shadow photographs, and from normal pressure measurements. Three Mach numbers of the un disturbed stream were covered, namely 0.4, 0.6, and 0.7. Estimates of the power absorbed by the com pressor, ignoring duct losses are made from meas urements of the mass of air sucked and the static pressure in the slots. Additional information was obtained on the adverse effect of a large radius of the forward lip of the slot, on the effect at Mo = 0.4 of an increase in slot width, and on the choke quantities for the slots. N26536 * Aeronautical Research Council (Gt. Brit.) THE EFFECT OF SLIPSTREAM ON THE LONGI TUDINAL STABILITY OF MULTIENGINED AIR CRAFT. D. E. Morris and J. C. Morrall. 1953. 9p. diagrs., tab. lARC R& M 2701; ARC 12.136. Formerly RAE Aero 2304) NACA RESEARCH ABSTRACTS NO.50 Flight measurements of longitudinal stability power off and poweron made on numerous aircraft have been analyzed and a generalized curve for estimating the contribution of slipstream to longitudinal stability. applicable to both flapsup and flapsdown cases, has been derived. Using this curve, the change in stability due to slipstream at a given value of CL can be estimated with a probable error of less than t 0.02 in the position of the neutral point. N26537 * Aeronautical Research Council (Gt. Brit.) LOWSPEED WINDTUNNEL TESTS ON TWO 45 DEG SWEPTBACK WINGS OF ASPECT RATIOS 4.5 AND 3.0 (MODELS A AND B). J. Trouncer and G. F. Moss. 1953. 43p. diagrs., 13 tabs. (ARC R & M 2710; ARC 10, 904. Formerly RAE Aero 2210) Lowspeed stability tests were made on two wings of aspect ratio 4.5 and 3.0. Both wings were of 450 sweepback, 4:1 taper ratio, and 14 percent thickness ratio. Tests included stability tests on the two wings without body or tail unit, tests with a body, fin and tailplane fitted and tests made with two types of nose flaps on the aspect ratio 4.5 wing. N26538* Aeronautical Research Council (Gt. Brit.) THE MEASUREMENT OF THE OVERALL DRAG OF AN AIRCRAFT AT HIGH MACH NUMBERS. D. J. Higton, R. H. Plascott and D. A. Clarke. 1953. 22p. diagrs., photo., tab. (ARC R & M 2748; ARC 11,429; ARC 12,237. Formerly RAE Aero 2241; RAE Aero 2309) This report describes the technique which has been developed to measure the overall drag of an air craft at high Mach numbers in both level flight and dives. It shows how improvements have been made both in flight and tunnel technique so that compari sons between fullscale and model tests have now become possible. Flight results from Meteor IV aircraft show close agreement between drag measured in level flight and in dives and later tests compare well with highspeed windtunnel meas urements on a 1/12th scale model. N26540* Aeronautical Research Council (Gt. Brit.) THE THEORETICAL PRESSURE DISTRIBUTIONS AROUND SOME CONVENTIONAL TURBINE BLADES IN CASCADE. T. J. Hargesi 1953. lOp. diagrs. (ARC R & M 2765; ARC 13,360. Formerly NGTE R. 67) By means of Relf's analogy between aerodynamic streamline flow and electric potential flow, the theoretical pressure distributions around a series of conventional turbine blades in cascade have been determined over a range of incidence covered in some previously reported aerodynamic tests. The theoretical pressure distributions and their variation with incidence provide the basis of an explanation of the observed aerodynamic performance. NACA RESEARCH ABSTRACTS NO.50 7 N26541* transition and separation. A brief consideration is also given to some of the effects of airstream turbu Aeronautical Research Council (Gt. Brit.) lence, weak shock waves from the walls, and heat A CALCULATION OF THE COMPLETE DOWNWASH transfer. IN THREE DIMENSIONS DUE TO A RECTANGULAR VORTEX. Doris E. Lehrian. 1953. 45p. diagrs., 15 labs. (ARC R&M 2771. Formerly ARC 11,786; S L. C 2252; Perf. 487; ARC 12, 221; S & C 2294; Perf. 537) A calculation of the complete downwash in three di mensions due to a rectangular vortex, is given for the limited range Z = +4. The downwash is com puted at selected positions, in planes normal to the plane of the vortex; these planes are spaced at even integral multiples of the semiwidth of the vortex, measured from the line of symmetry. Values are tabulated for Z in the range (0, 4) and a set of graphs is also included for 0 < Z < 2; they are to be used in conjunction with the "Tables of Complete Downwash due to a Rectangular Vortex" (R. & M. 2461). MISCELLANEOUS NACA TN 2991 Errata No. 1 on "ACCELERATIONS AND PASSEN GER HARNESS LOADS MEASURED IN FULL SCALE LIGHTAIRPLANE CRASHES." A. Martin Elband, Scott H. Simpkinson and Dugald 0. Black. August 1953. UNPUBLISHED PAPERS N20019A* REVIEW OF PUBLISHED DATA ON THE EFFECT OF ROUGHNESS ON TRANSITION FROM LAMINAR TO TURBULENT FLOW. Hugh L. Dryden. 6p. diagrs. (Reprint from Journal of the Aeronautical Sciences, v. 20, no. 7, July 1953, p. 477482). A review is presented of the published data on the effect of roughness, especially single roughness elements, on transition from laminar to turbulent flow, in which an attempt is made to reanalyze and correlate the available information. The paper also discusses available data on the effect of distributed roughness on transition on a flat plate, as well as some of the published data on roughness effects on transition on airfoils. N25045* SOME FACTORS CONTRIBUTING TO SCALE EFFECT AT SUPERSONIC SPEEDS. Ira H. Abbott. 34p. diagrs., photos. (Presented at the fourth meeting of the Advisory Group for Aeronautical Research and Development, Wind Tunnel and Model Testing Panel, London, September 311, 1953) This paper presents information on the Reynolds number effects at supersonic speeds on skin friction, NACALanldev 10953 4M UNIVERSITY OF FLORIDA 3 1262 08153 274 8 
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