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National Advisory Committee for Aeronautics Research Abstracts N0.27 AUGUST 15,1952 CURRENT NACA REPORTS NACA Rept. 1025 EXPERIMENTAL AND THEORETICAL STUDIES OF AREA SUCTION FOR THE CONTROL OF THE LAMINAR BOUNDARY LAYER ON AN NACA 64A010 AIRFOIL. Albert L. Braslow, Dale L. Burrows, Neal Tetervin and Fioravante Visconti. 1951. ii, 19p. diagrs., photos. (NACA Rept. 1025. Formerly TN 1905; TN 2112) Lowturbulence windtunnel tests were made of an NACA 64A010 airfoil section having a porous surface to investigate continuous suction as a means of con trolling the laminar boundary layer. Fullchord lam inar flow and large net reductions in drag were ob tained up to a Reynolds number of approximately 20 x 10. It seems likely from the experimental re sults and a related theoretical analysis that attain ment of fullchord laminar flow by means of continu ous suction through a porous surface will be possible at any value of the Reynolds number provided that the airfoil surfaces are maintained sufficiently smooth and fair and provided that outflow of air through the surface is prevented. NACA Rept. 1029 COMPRESSIVE STRENGTH OF FLANGES. Elbridge Z. Stowell. 1951. 14p. diagrs., tab. (NACA Rept. 1029. Formerly TN 2020) The maximum compressive stress carried by a hinged flange is computed from a deformation theory of plasticity combined with the theory for finite de flections for this structure. The computed stresses agree well with those found experimentally. Empiri cal observation indicates that the results will also apply fairly well to the more commonly used flanges which are not hinged. NACA Rept. 1030 INVESTIGATION OF SEPARATION OF THE TURBU LENT BOUNDARY LAYER. G. B. Schubauer and P. S. Klebanoff, National Bureau of Standards. 1951. 20p. diagrs., photos., 8 tabs. (NACA Rept. 1030. Formerly TN 2133) An investigation was conducted on a turbulent bound ary layer near a smooth surface with pressure gradients sufficient to cause flow separation. The Reynolds number was high, but speeds were entirely within the incompressible flow range. The investiga tion consisted of measurements of mean flow, three components of turbulence intensity, turbulent shear ing A.tres., and correlati$ns between two fluctuation components at a point and between the same conmpi nenk!t dilterent point.. Results are given in the form ol tables, and graphs.. The discussion deal. first with separation and then with the more funda mental question of basic concepts of turbulent flow. NACA Rept. 1035 ANALYSIS OF MEANS OF IMPROVING THE UN CONTROLLED LATERAL MOTIONS OF PERSONAL AIRPLANES. Marion 0. McKinney, Jr. 1951. ii, 9p. diagrs., 3 tabs. (NACA Rept. 1035. Formerly TN 1997) A theoretical analysis of means of improving the un controlled lateral motions of personal airplanes is presented. The purpose of this analysis was to deter mine whether such airplanes could be made capable of flying uncontrolled for an indefinite period of time without getting into a dangerous attitude and for a reasonable period of time without deviating ex cessively from the original course. NACA Rept. 1037 GENERAL METHOD AND THERMODYNAMIC TABLES FOR COMPUTATION OF EQUILIBRIUM COMPOSITION AND TEMPERATURE OF CHEMICAL REACTIONS. Vearl N. Huff, Sanford Gordon and Virginia E. Morrell. 1951. ii, 57p., 45 tabs. (NACA Rept. 1037. Formerly NACA TN 2113, TN 2161) A rapidly convergent successive approximation process is described that simultaneously determines both composition and temperature resulting from a chemical reaction. This method is suitable for use with any set of reactants over the complete range of mixture ratios as long as the products of reaction are ideal gases. An approximate treatment of limited amounts of liquids and solids is also included. This method is particularly suited to problems having a large number of products of reaction and to problems that require determination of such properties as specific heat or velocity of sound of a dissociating mixture. The method presented is applicable to a wide variety of problems that include (1) combustion at constant pressure or volume; and (2) isentropic expansion to an assigned pressure, temperature, or Mach number. Tables of thermodynamic functions needed with this method are included for 42 sub stances for convenience in numerical computations. *AVAILABLE ON LOAN ONLY. ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1724 F ST., NW., WASHINGTON 25, D. C, CITING CODE NUMBER ABOVE EACH TITLE, THE REPORT TITLE AND AUTHOR. 2 NACA Rept. 1038 WINDTUNNEL INVESTIGATION OF AIR INLET AND OUTLET OPENINGS ON A STREAMLINE BODY. John V. Becker. 1951. ii, 21p. diagrs., photos., 3 tabs. (NACA Rept. 1038. Formerly ACR, November 1940) Air inlet openings at the nose of a streamline body and outlets at the tail and at the 21percent and 63 percent stations of the body were investigated with the objective of developing lowdrag highcritical speed configurations. Drag, pressuredistribution, and boundarylayer measurements were made. Inlet and outlet opening shapes were found which caused no increase in drag over that of the basic streamline body to which the openings were applied. The criti cal Mach number of the ducted body was also equal to that of the basic body over a wide range of internal air flow rates. NACA Rept. 1042 SOME EFFECTS OF NONLINEAR VARIATION IN THE DIRECTIONALSTABILITY AND DAMPINGIN YAWING DERIVATIVES ON THE LATERAL STA BILITY OF AN AIRPLANE. Leonard Sternfield. 1951. ii, 9p. diagrs., tab. (NACA Rept. 1042. Formerly TN 2233) The effect of nonlinear stability derivatives on the lateral stability of an airplane is analyzed. Motions are calculated on the assumption that the values of the directionalstability derivative and the damping inyawing derivative are functions of the angle of sideslip. The application of the Laplace transform to the calculation of an airplane motion when certain types of nonlinear derivatives are present is de scribed in detail. NACA Rept. 1047 THE STABILITY OF THE COMPRESSION COVER OF BOX BEAMS STIFFENED BY POSTS. Paul Seide and Paul F. Barrett. 1951. ii, 16p. diagrs., 3 tabs. (NACA Rept. 1047. Formerly TN 2153) An investigation is made of the buckling of the com pression cover of poststiffened box beams subjected to end moments. Charts are presented for the deter mination of the minimum post axial stiffnesses and the corresponding compressive buckling loads re quired for the compression cover to buckle with nodes through the posts. Application of the charts to de sign and analysis and the limitations of their use are discussed. NACA Rept. 1049 EXPERIMENTAL INVESTIGATION OF THE EFFECT OF VERTICALTAIL SIZE AND LENGTH AND OF FUSELAGE SHAPE AND LENGTH ON THE STATIC LATERAL STABILITY CHARACTERISTICS OF A MODEL WITH 45 SWEPTBACK WING AND TAIL SURFACES. M. J. Queijo and Walter D. Wolhart. 1951. ii, 29p. diagrs., photos., 4 tabs. (NACA Rept. 1049. Formerly TN 2168) NACA RESEARCH ABSTRACTS NO.27 Results are presented of an investigation to deter mine the effects of verticaltail size and length and of fuselage shape and length on the lateralstatic stability derivatives of a midwing airplane model with 45 sweptback wing and tail surfaces. Inter ference between the various model components also is evaluated. NACA Rept. 1052 A SUMMARY OF LATERALSTABILITY DERIVA TIVES CALCULATED FOR WING PLAN FORMS IN SUPERSONIC FLOW. Arthur L. Jones and Alberta Alksne. 1951. ii, 35p. diagrs., 3 tabs. (NACA Rept. 1052) Values of the lateralstability derivatives for wings at supersonic speeds, calculated using the linearized theory for compressible flow, are presented in the form of design charts showing the variations of the derivatives with Mach number and aspect ratio for six plan forms. Limitations in the applicability and availability of the lateralstability derivatives are discussed. NACA Rept. 1059 A BIHARMONIC RELAXATION METHOD FOR CAL CULATING THERMAL STRESS IN COOLED IRREGU ULAR CYLINDERS. Arthur G. Holms. 1952. ii, 19p. diagrs., 4 tabs. (NACA Rept. 1059. Formerly TN 2434) A numerical method was developed for calculating thermal stresses in irregular cylinders cooled by one or more internal passages. The use of relaxation methods and elementary methods of finite differences was found to give approximations to the correct values when compared with previously known solutions for concentric circular cylinders possessing symmetrical and asymmetrical temperature distributions. NACA Rept. 1060 DETAILED COMPUTATIONAL PROCEDURE FOR DESIGN OF CASCADE BLADES WITH PRESCRIBED VELOCITY DISTRIBUTIONS IN COMPRESSIBLE PO TENTIAL FLOWS. George R. Costello, Robert L. Cummings and John T. Sinnette, Jr. 1952. ii, 14p. diagrs., 9 tabs. (NACA Rept. 1060. Formerly TN 2281) A detailed stepbystep computational outline is pre sented for the design of cascade blades having a pre scribed velocity distribution on the blade in a poten tial flow of the usual compressible fluid. This outline is based on the assumption that the magnitude of the velocity in the flow of the usual compressible non viscous fluid is proportional to the magnitude of the velocity in the flow of a compressible nonviscous fluid with linear pressurevolume relation. The computa tional procedure includes several ways of adjusting the prescribed velocity to satisfy the restriction im posed by the method. Tables of coefficients are given for evaluating the necessary integrals, includ ing the determination of the harmonic conjugate. Nu merical examples are included. NACA RESEARCH ABSTRACTS N0.27 NACA Rept. 1061 EFFECT OF INITIAL MIXTURE TEMPERATURE ON FLAME SPEED OF METHANEAIR, PROPANE AIR AND ETHYLENEAIR MIXTURES. Gordon L. Dugger. 1952. ii, 12p. diagrs., photo., 3 tabs. (NACA Rept. 1061. Formerly TN 2170; TN 2374) Flame speeds based on the outer edge of the shadow cast by the laminar Bunsen cone were determined as functions of composition for methaneair mixtures at initial mixture temperatures ranging from 132 to 342 C and for propaneair and ethyleneair mixtures at initial mixture temperatures ranging from 73 to 3440 C. The data showed that maximum flame speed increased with temperature at an increasing rate. The percentage change in flame speed with change in initial temperature for the three fuels followed the de creasing order, methane, propane, and ethylene. Empirical equations were determined for maximum flame speed as a function of initial temperature over the temperature range covered for each fuel. For each fuel it was found that, with a fixed parallelbeam shadowgraph system, the ratio of flame speed based on the outer edge of the shadow cast by the flame cone to flame speed based on the inner edge of the shadow was a constant, independent of temperature or com position. The flame speed of propaneair flames was independent of tube diameter from 10 to 22 milli meters or streamflow Reynolds number from 1500 to 2100. The observed effect of temperature on flame speed for each of the fuels was reasonably well pre dicted by either the thermal theory as presented by Semenov or the squareroot law of Tanford and Pease. The importance of active radicals in flame propaga tion was indicated by a simple linear relation between maximum flame speed and equilibrium radical con centrations for all three fuels. NACA TN 2643 SPAN LOAD DISTRIBUTIONS RESULTING FROM ANGLE OF ATTACK, ROLLING, AND PITCHING FOR TAPERED SWEPTBACK WINGS WITH STREAM WISE TIPS. SUPERSONIC LEADING AND TRAILING EDGES. John C. Martin and Isabella Jeffreys. July 1952. 143p. diagrs., 6 tabs. (NACA TN 2643) On the basis of the linearized supersonicflow theory the span load distributions resulting from constant angle of attack, from steady rolling, and from steady pitching were calculated for a series of thin swept back tapered wings with streamwise tips and with supersonic leading and trailing edges. The results are valid for the Mach number range for which the Mach line from either wing tip does not intersect the remote halfwing. The results of the analysis are presented as a series of design charts. Some illus trative variations of the spanwise distributions of circulation with the various design parameters are also presented. NACA TN 2656 A BLADEELEMENT ANALYSIS FOR LIFTING ROTORS THAT IS APPLICABLE FOR LARGE IN FLOW AND BLADE ANGLES AND ANY REASON 3 ABLE BLADE GEOMETRY. Walter Castles, Jr. and Noah C. New, Georgia Institute of Technology. July 1952. 63p. diagrs., 7 tabs. (NACA TN 2656) A bladeelement analysis for lifting rotors is pre sented which is applicable for large inflow and blade angles and any reasonable blade geometry and should therefore be useful for convertaplane as well as heli copter calculations. Simple approximate relations between rotor thrust and flightpath velocity compo nents and rotor blade angles, torque, and inplane forces are derived and these solutions, based on the assumption of a triangular distribution of blade circulation and a parabolic variation of profile drag with lift, are sufficiently accurate for preliminary calculations and determination of equilibrium angle of attack and lateral tilt of the tippath plane. More exact bladeelement equations are then derived for the relations between thrust and flightpath velocity components and equilibrium blade angles, torque, and inplane forces and moments. NACA TN 2675 MEASUREMENTS OF FLYING QUALITIES OF AN F47D30 AIRPLANE TO DETERMINE LATERAL AND DIRECTIONAL STABILITY AND CONTROL CHARACTERISTICS. R. Fabian Goranson and Christopher C. Kraft, Jr. July 1952. 61p. diagrs., photos., 2 tabs. (NACA TN 2675) Tests have been made of the flying qualities of an F47D30 airplane to determine the lateral and directional stability and control characteristics. Data are also presented of the aileron hingemoment characteristics. NACA TN 2715 THE THEORETICAL CHARACTERISTICS OF TRIANGULARTIP CONTROL SURFACES AT SUPER SONIC SPEEDS. MACH LINES BEHIND TRAILING EDGES. Julian H. Kainer and Mary Dowd King. July 1952. 76p. diagrs., 4 tabs. (NACA TN 2715) By means of linearized theory, generalized expres sions in closed form have been obtained for the char acteristics due to controlsurface deflection (CL6 C16, Cm6, and Ch6 ) and due to wing angle of attack (Cha) for wing plan forms having triangular tip control surfaces at supersonic speeds. The analysis considers wing trailingedge sweep, control surface geometry, and Mach number for the deflec tion characteristics. For Cha, the effects of wing leadingedge sweep and aspect ratio are also included. The analysis is limited to configurations where the trailing edges are supersonic and where the inner most Mach lines from the leading edge of the control surface root chord do not intersect the wing root chord. 4 NACA TN 2737 PLASTIC STRESSSTRAIN RELATIONS FOR COM BINED TENSION AND COMPRESSION. Joseph Marin and H. A. B. Wiseman, Pennsylvania State College. July 1952. 61p. diagrs., photos., 2 tabs. (NACA TN 2737) I'lastic stressstrain relations for biaxial tension compression principal stresses were determined for a 14ST4 aluminum alloy. Constantstressratio tests provided control data and information on the in fluence of biaxial stresses on the yield strength of the material. Variablestressratio tests were made to determine whether the deformation or flow theory a.ii ee better with actual plastic stressstrain rela tions and special tests were conducted to check the validity of various assumptions made in these theories. NACA TN 2739 NUMERICAL DETERMINATION OF INDICIAL LIFT AND MOMENT FUNCTIONS FORA TWO DIM ENSIONAL SINKING AND PITCHING AIRFOIL AT MACH NUMBERS 0.5 AND 0.6. Bernard Mazelsky and Joseph A. Drischler. July 1952. 37p. diagrs., 4 tabs. (NACA TN 2739) Approximate values for the indicial lift and moment functions on a sinking and pitching airfoil are calcu lated for Mach numbers 0.5 and 0.6. The indicial lift function associated with an airfoil penetrating a sharpedge gust is also determined approximately at Mach numbers 0.5, 0.6, and 0.7. NACA TN 2741 INVESTIGATION OF THE INFLUENCE OF FUSE LAGE AND TAIL SURFACES ON LOWSPEED STATIC STABILITY AND ROLLING CHARACTERIS TICS OF A SWEPTWING MODEL. John D. Bird, Jacob H. Lichtenstein and Byron M. Jaquet. July 1952. 18p. diagrs., photo. (NACA TN 2741. For merly RM L7H15) Results are presented of a windtunnel investigation to determine influence of the fuselage and tail on static stability and rotary derivatives in roll of a model having 45 sweptback wing and tail surfaces. The wing alone and the model without the horizontal tail showed marginal longitudinal stability near maxi mum lift. The longitudinal stability of the complete model was satisfactory. The vertical tail produced larger increments of rate of change of lateralforce and yawingmoment coefficients with wingtip helix angle than the fuselage or the horizontal tail. NACA TN 2744 PRACTICAL CALCULATION OF SECONDORDER SUPERSONIC FLOW PAST NONLIFTING BODIES OF REVOLUTION. Milton D. Van Dyke. July 1952. 62p. diagrs., 2 tabs., 2 charts. (NACA TN 2744) NACA RESEARCH ABSTRACTS NO. 27 Calculation of secondorder supersonic flow past bodies of revolution at zero angle of attack is de scribed in detail, and reduced to routine computation. Use of an approximate tangency condition is shown to increase the accuracy for bodies with corners. Tables of basic functions and standard computing forms are presented. The procedure is summarized so that one can apply it without necessarily under standing the details of the theory. A sample calcula tion is given, and several examples are compared with solutions calculated by the method of character istics. NACA TN 2745 INFLUENCE OF CHEMICAL COMPOSITION ON RUPTURE TEST PROPERTIES AT 1500 F OF FORGED CHROMIUMCOBALT NICKELIRON BASE ALLOYS. J. W. Freeman, J. F. Ewing and A. E. White, University of Michigan. July 1952. 69p. diagrs., photos., 2 tabs. (NACA TN 2745) Six1iytwo alloys involving systematic individual varia tions of 10 elements present in a forged chromium cobaltnickeliron base alloy in the solutiontreated and aged condition and simultaneous variation of molybdenum, tungsten, and columbium were rupture tested at 1500 F. It was found that the elements can be varied individually between quite wide limits with out significantly changing the rupture properties. The results were similar to those previously obtained at 1200 F (NACA Rept. 1058) except that increased chromium was not beneficial at 1500 F and the saturation effect for molybdenum and tungsten did not occur at 1200 F. NACA TN 2746 PREVIEW OF BEHAVIOR OF GRAIN BOUNDARIES IN CREEP OF ALUMINUM BICRYSTALS. F. N. Rhines and A. W. Cochardt, Carnegie Institute of Technology. July 1952. 40p. diagrs., photos. (NACA TN 2746) Gliding of aluminum bicrystals along their mutual boundary was studied during creep tests at 200 to 650 C and 1 to 100 psi. The direction of motion de pends only upon the direction of maximum resolved shear stress in the boundary plane. The magnitude of the motion increases with magnitude of resolved shear stress, temperature, and degree of mismatch between orientations of the crystals. Gliding rate is cyclic and for each temperature there is a stress below which gliding does not occur. At low temperature and small orientation difference an induction period pre cedes gliding. Gliding operates in a relatively thick zone of crystalline metal on both sides of the bound ary. Portions of the grains away from the boundary degenerate into blocks that move with respect to each other, parallel to the tension axis, and that rotate about octahedral axes of the original crystal. NACA RESEARCH ABSTRACTS NO.27 NACA TN 2750 MATRIX AND RELAXATION SOLUTIONS THAT DETERMINE SUBSONIC THROUGH FLOW IN AN AXIALFLOW GAS TURBINE. ChungHua Wu. July 1952. 65p. diagrs., 7 tabs. (NACA TN 2750) A method recently developed for determining the steady flow of a nonviscous compressible fluid along a relative stream surface between adjacent blades in a turbomachine was applied to investigate subsonic through flow in a singlestage axialflow gas turbine. For both incompressible and compressible flows, the variation in flow on the stream surface was deter mined by using both the relaxation and the matrix methods. In all solutions considered, convergence was obtained without difficulty. The compressibility of the gas and the radial twist of the stream surface had equally significant effects on the considerable amount of radial flow obtained. The results of these accurate calculations provide a basis for evaluation of simpler, more approximate methods for computing subsonic through flow in turbomachines. NACA TN 2754 A METHOD OF SELECTING THE THICKNESS, HOLLOWNESS, AND SIZE OF A SUPERSONIC WING FOR LEAST DRAG AND SUFFICIENT BENDING STRENGTH AT SPECIFIED FLIGHT CONDITIONS. James L. Amick. July 1952. 38p. diagrs. (NACA TN 2754) The method described, which is suitable for use in preliminary missile design, is based on simple beam theory and takes into account the effects of wing weight and type of skin thickness distribution. For cases in which the wing weight is negligible, a simplification of the general method gives the best liftdrag ratio consistent with the bending strength requirement of a given combination of plan form and profile shape at a given Mach number as a function of a single design parameter. An example of the appli cation of the method to a diamond wing at Mach num ber 2.0, for a range of specified flight conditions, is presented. NACA RM E52C31 TRUE AIRSPEED MEASUREMENT BY IONIZATION TRACER TECHNIQUE. Bemrose Boyd, Robert G. Dorsch and George H. Brodie. July 1952. 37p. diagrs., photos. (NACA RM E52C31) Ion bundles produced in a pulseexcited corona dis charge are used as tracers with a radarlike pulse transittime measuring instrument in order to pro vide a measurement of airspeed that is independent of all variables except time and distance. The resulting instrumentation need not project into the air stream and, therefore, will not cause any interference in supersonic flow. The instrument was tested at Mach numbers ranging from 0.3 to 3.8. Use of the proper instrumentation and technique results in accuracy of the order of 1 percent. 5 NACA RM E52E12 USE OF CHOKED NOZZLE TECHNIQUE AND EXHAUST JET DIFFUSER FOR EXTENDING OPER ABLE RANGE OF JETENGINE RESEARCH FACIL ITIES. John H. Povolny. July 1952. 17p. diagrs. (NACA RM E52E12) An investigation has been conducted to determine the increase in the useful ranges of flight conditions that may be obtained with a given jetengine research facility when the choked nozzle technique or the exhaust jet diffuser, or both, are employed. This report describes these two methods, presents the considerations involved in their application, and gives typical results of their use as well as confir mation of the accuracy of data obtained by utilization of these techniques. The validity and accuracy of the choked nozzle technique and the associated area  pressuredifferential thrust correction term were substantiated by turbojetengine and exhaustnozzle performance data covering a range of nozzle pres sure ratios up to about 10. It was demonstrated by calculations for a typical turbojet engine installed in a typical altitude test facility that a considerable in crease in the range of flight conditions that can be in vestigated may be obtained by use of the choked nozzle technique. It was also demonstrated that the range of facility exhaust pressures or exhauster flows may be increased by use of the exhaust jet diffuser. BRITISH REPORTS N16326* Aeronautical Research Council (Gt. Brit.) PUBLISHED REPORTS AND MEMORANDA OF THE AERONAUTICAL RESEARCH COUNCIL. 1952. 5p. (ARC R & M 2350) Lists R & M's 2251 through 2350 and gives the Coun cil number, the title and the author for each. N16327* Aeronautical Research Council (Gt. Brit.) TUNNEL BLOCKAGE NEAR THE CHOKING CONDI TION. A. Thorn and Myra Jones. 1952. 16p. diagrs., 2 tabs. (ARC R & M 2385; ARC 8606; ARC 8878. Formerly RAE Aero 2020; RAE Aero 2056; ARC 9854; FM 968) By a consideration of mass flow, blockage correc tions are estimated up to the choking speed of the tun nel, and a curve is given showing the blockage for a Rankine oval in terms of the uncorrected choking Mach number. The theory of the method of deter mining blockage from wall pressures is examined more critically for the twodimensional cases and extended to certain threedimensional cases. 6 N16328* Aeronautical Research Council (Gt. Brit.) BOUND AND TRAILING VORTICES IN THE LINE ARISED TH Oi'Y OF SUPERSONIC FLOW, AND THE DOWNWASH IN THE WAKE OF A DELTA WING. A. Robinson and J. H. HunterTod. 1952. 14p. dij.rs. (ARC R :. M 2409; ARC 11,296. Formerly College of Aeronautics, Crfijiilid Rept. 10) The field of flow round a flat airfoil at incidence can be regarded in linearized theory as the result of both bound and trailing vortices for supersonic as well as for lowspeed flight. This leads to a convenient method, given the lift distribution over an iirfil. for ( .ili ulailinL the flow round it at supersonic speeds. As an application of the results, the downwash is cal culated in the wake of a delta wing lying within the Mach cone .iimnatin4 from its apex. The downwash is found to be least just aft the trailing edge and is everywhere less than the downflow at the airfoil. It increases steadily to a lniiliii value which is attain ed virtually within two chord lengths of the trailing edil't The ratio of the downwash at any point in the wake to the downflow at the airfoil decreases with in creasing Mach number and apex angle. N16329* Aeronautical Research Council (Gt. Brit.) THE AERODYNAMIC DERIVATIVES WITH RESPECT TO SIDESLIP FOR A DELTA WING WITH SMALL DIHEDRAL AT ZERO INCIDENCE AT SUPERSONIC SPEEDS. A. Robinson and J. H. HunterTod. 1952. 14p. diagrs. (ARC R M 2410; ARC 11,322. For merly College of Aeronautics, Cranfield. Rept. 12) Expressions are derived for the sideslip derivatives on the assumptions of the linearized theory of flow for a delta wing with small dihedral flying at super sonic speeds. A discussion is included in the Appen dix on the relation between two methods that have been evolved for the treatment of aerodynamic force problems of the delta wing lying within its apex Mach cone. When the leading edges are within the Mach cone from the apex, the pressure distribution and the rolling moment are independent of Mach number but dependent on aspect ratio. When the leading edges are outside the apex Mach cone, the nondimensional rolling derivative is, in contrast to the other case, dependent on Mach number and independent of aspect ratio; the other derivatives and the pressure, how ever, are dependent on both variables. N16330" Aeronautical Research Council (Gt. Brit.) USE OF NEGATIVE CAMBER IN THE TRANSONIC SPEED RANGE. W. F. Hilton. 1952. 5p. diagrs. (ARC R & M 2460. Formerly ARC 10, 421; FM 1077; S & C 2100) Certain difficulties are experienced when attempting free flight in the transonic speed range (0. 8 to 1. 2 of the speed of sound). These difficulties fall into two main classes: namely, (1) overcoming the somewhat high air resistance by means of improvements in en NACA RESEARCH ABSTRACTS N0.27 gine design, and (2) balancing the aircraft for stable horizontal flight. The latter problem is considered in this paper. Changes of trim are caused by sudden loss of wing lift in the transonic range, which de creases the downwash over the tail, and possibly re sults in an uncontrollable noseheavy dive. The use of negative wing camber for minimizing these effects is suggested. and the suggestion is found to be sup ported by windtunnel experiments. N16331* Aeronautical Research Council (Gt. Brit.) EXPERIMENTS ON THIN TURNING VANES. C. Salter. 1952. 29p. diagrs., photo., 9 tabs. (ARC R & M 2469. Formerly ARC 10, 039; TP 164) The experiments recorded in this paper refer to a variety of tests made on corner vanes (mainly thin curved plates) in a tunnel 1 foot square, with the ob ject of indicating whether and to what extent they would be suitable for deflecting the air stream around the corners of returnflow type wind tunnels. It was found that the air stream could be turned very satis fa( t'rilly by the use of thin vanes and that the loss of head in a 90 corner was not as great as had been ex pected. Taking the pitch of the vanes to be the geo metrical pitch in a direction perpendicular to the axis of the tunnel and the chord to be the distance between the centers of radii of the leading and trailing edges, the optimum ratio of pitch to chord for minimum loss was shown to be about 0. 25. At that P/C ratio and Reynolds number of 2 x 105, based on chord length and air velocity of approach, the loss of total head in the corner was found to be 0. 12 of the dynamic head, while in the region outside the boundary layer the minimum value was only 0. 06. To ensure stability of the stream, a P/C ratio of 0. 2 is recommended and it appears that, in order to maintain a thin bound ary layer, the actual number of vanes in any one cor ner should be not less than say 15 to 20. A clean en try of the streamlines into the vane unit can be estab lished by giving the leadingedge tangents an incidence of about 5 relative to the upstream axis of the tunnel. The minimum corner loss with 45 deflection was found to be about 0.05 of the dynamic head with a P/C ratio of roughly 0. 4. A set of "Collar" vanes with splitters gave very satisfactory turning of the stream with a loss of head in the corner of 0. 17 of the dynamic head, at a Reynolds number of 150, 000, the corresponding figure in the region outside the bound ary layer being 0.12. As the effect of the walls in assisting the deflection of the stream is so large it is suggestedthat the staticpressure distribution is as good a criterion as any of the effectiveness of the cor ner. Uniformity of static pressure over both up stream and downstream crosssection to within a short distance from the walls is a good test for the correct shape and setting of the vanes. N16332* Aeronautical Research Council (Gt. Brit.) STRAIN GAUGE FLUTTER TESTS ON A 4BLADE PROPELLER WITH DURALUMIN BLADES. J. Kettlewell and H. G. Ewing. 1952. 4p. diagrs. (ARC R & M 2471; ARC 9876. Formerly RAE Tech. Note SME 354) NACA RESEARCH ABSTRACTS N0.27 This note describes tests made on a fluttering duralu min propeller with strain gages fitted. Three dif ferent blade settings in the stalling region were in vestigated. The flutter encountered was purely tor sional; and the amplitude of strain recorded was at no time serious. N16333* Aeronautical Research Council (Gt. Brit.) WINDTUNNEL TESTS ON THE 30 PER CENT. SYMMETRICAL GRIFFITH AEROFOIL WITH EJEC TION OF AIR AT THE SLOTS. N. Gregory, W. S. Walker and W. G. Raymer. 1952. lip. diagrs. (ARC R & M 2475. Formerly ARC 10, 097; FM 1018, Perf. 251) It has been shown by Preston (1946) that ejection of air at the point of velocity discontinuity on a 16.2 percent thick Griffith suction airfoil prevents separation, and that if sufficient air is ejected, the drag is reduced. The present tests were undertaken to apply this prin ciple to the 30 percent Griffith airfoil and to investi gate the effect on lift by pressureplotting the airfoil. Ejection of air was found to prevent separation, but about 66 percent more air was required than with suc tion. Three times the suction quantity of air, when ejected, reduced the drag to the low values associated with suction. At R = 0. 96 million, the range of the tests was 018 incidence and 014 flap angle. At 180 incidence and 14 flap angle, a CNF of 2. 5 was obtained, giving approximately the same liftcurve slope as with suction. Above this angle of incidence, the pump capacity was not large enough for unsepa rated flow to be attained. With separation prevented, the pitching moments were the same as with suction, but the hinge moments were sensitive to small changes of blowing quantity. At R = 2. 88 million, the pump capacity was insufficient to prevent a partial stall at 6 incidence as occurred with suction. Curves of CNF, CQ, CM, CH, CD and velocity dis tribution when blowing are given, and comparisons are made with corresponding curves obtained with suction and with no suction. The same lift and pitch ing moments are obtained at any incidence with blow ing and with suction, but the suction quantities are about 40 percent less than the blowing quantities. The hinge moments are greatly different with blowing, and increase of the normal force. N16334* Aeronautical Research Council (Gt. Brit.) THE EFFECT OF POINTED TIPS ON WING LOAD ING CALCULATIONS. V. M. Falkner. 1952. 4p. diagr., tab. (ARC R & M 2483. Formerly ARC 10, 037; S & C 2063; Perf. 235) The standard formulas for calculation of wing loading by vortex lattice theory, which involve a sequence starting with 4f( p2), are not strictly valid for pointed wing tips. On the other hand, the use of a sequence starting with 1 12 leads to overcorrection. By considering the mean of these two solutions for a delta wing, it is shown that the error introduced by the use of the standard sequence is small and is on the safe side as regards bending moments. 7 N16335* Aeronautical Research Council (Gt. Brit.) FLIGHT MEASUREMENTS OF THE PRESSURE DIS TRIBUTION ON A TEMPEST WING UP TO A MACH NUMBER OF 0. 8. K. Eyre. 1952. 15p. diagrs. (ARC R & M 2489; ARC 10, 550. Formerly Hawker Aircraft, Ltd. Design Rept. 1131) This investigation was to determine the effect of Mach number on the pressure distribution of the Hawker airfoil section as fitted to Tempest and Fury aircraft. The distributions show the typical super sonic flow conditions obtained over an airfoil of this type. The airfoil does not show any unusual charac teristics and there are no large changes in pitching moment and position of the aerodynamic center below a Mach number of about 0. 75. In all cases the favor able pressure gradient in front of the shock wave is fairly gradual, but changes rapidly at the shock wave. There is, however, no sign of a breakdown in flow at the trailing edge, where the distributions are similar for all values of Mach number. The values obtained for critical Mach number agree quite well with those given in the Royal Aeronautical Society data sheets. The variations of centerofpressure position, pitch ing moment, and aerodynamiccenter position are in quite good agreement with German windtunnel tests made on an approximately similar airfoil. N16336* Aeronautical Research Council (Gt. Brit.) 24FT. WINDTUNNEL TESTS ON A "PADDLE BLADE" PROPELLER. A. B. Haines. 1952. 15p. diagrs. (ARC R & M 2403; ARC 10, 308. Formerly RAE Aero 2172) The report contains the results of tests made in the Royal Aircraft Establishment 24foot tunnel on a de Havilland propeller having conventional Clark Y sec tions and the paddleblade type of plan form. The report describes analyses of the results made both by singleradius and fullstrip theory methods. The data provided will enable estimates to be made of the performance of paddleblade propellers under various flight conditions and, in particular, will give a quan titative idea of the improvements to be expected for high power loadings, that is, beyond the stall. The principal general conclusions are: (1) Use of the same lift/drag data in eight radius strip theory calcu lations for both paddleblade and normal propellers is justified. (2) The gains found in flight tests for paddleblade propellers must be principally ascribed to the effects of NACA 16 series sections. (3) RAE charts based on singleradius data may be pessimis tic in predicting the peak efficiencies of paddle pro pellers at high tip speed (by 1. 5 to 2 percent for MT = 0. 85). (4) The mean stalling performance is better than would be expected even on an activity fac tor basis. This enables other design modifications, for example, thin roots, 16 series sections, etc. to be made to improve the top speed performance with out causing serious losses at takeoff or climb. (5) Therefore, the general conclusion is that use of the paddle type of plan form is fully justifiable even though it does not appear to offer any direct shock stalling relief. 8 N16337 Aeronautical Research Council (Git. Brit.) TESTS ON THREE EQUILATERAL TRIANGULAR PLATES IN THE COMPRESSED AIR TUNNEL. R. Jones and C. J. W. Miles. 1952. 7p. diagrs., 4 tabs. (ARC R M 2518. Formerly ARC 9973; Perf. 222) The report gives CL, CD, and Cm on three equi lateral triangular plates with sides 26, 36, and 47. 8 inches. The object of the tests was to estimate the size of sweptback wings of that form, which can be tested in the compressed air tunnel without the application of excessive tunnel corrections. After applying the usual tunnel corrections and making small corrections due to dissymmetry in the models, it was found that CL on the three models .,.., i:ed if the results on the 36inch and 47. 8inch models were multiplied by 1. 01 and 1. 05, respectively. Fair .ILt*i t liiCl was obtained on CD for the three models. The mean position of the center of pressure is about 0.45 of the height of the tri.incl,., forward ot the trail ing edge. N16338* Aeronautical Research Council (Gt. Brit.) THE YAWING VIBR'ATIONS OF AN AIRCRAFT. J. Morris and G. S. Green. 1952. llp. diagrs. (ARC R ' M 2525; ARC 9482. Formerly RAE SME 4027) This report gives a theoretical method for calculating the natural frequencies and modes of yawing vibration of a complete aircraft. The basic feature of the treatment is the replacement of the continuous mass system by one consisting of a finite number of dis crete masses elastically interconnected. In the course of the analysis, use is made of the deflection coefficient artifice in the formation of the equations of motion, and the escalator process in their mar shalling and numerical solution. The method has been applied to a singleengined fighter aircraft, for which the results of a resonance test were available. These results appear to be some 40 percent in excess of their calculated counterparts and no satisfactory explanation occurs to the authors to account for this incompatability. N16339* Aeronautical Research Council (Gt. Brit.) AIRCRAFT LANDINu GEAR: GROUND LOADS WHEN SPINNINGUP THE WHEELS AT TOUCH DOWN. J. W. Blinkhorn. 1952. 13p. diagrs., 2 tabs. (ARC R z M 2588; ARC 11, 627. Formerly RAE Tech. Note Mech. Eng. 16) The investigation covers all combinations of landing speed, coefficient of friction between the tire and the ground, and harshness of landing, for any type of pneumatic tire and wheel unit. Particular atten tion has been given to landing speeds between 50 and 150 mph, coefficients of friction from 0 to 2. 0, and landings giving vertical wheel accelerations of ig, NACA RESEARCH ABSTRACTS NO.27 2g, 3g, and 4g. It was found that for any landing, the vertical reaction, at any wheel which has just finished spinning up increases with increase in the moment of inertia of the wheel and tire unit, and the landing speed, and decreases with increase in the free tire radius, the aircraft weight, the time to reach the maximum vertical wheel reaction, and the coefficient of friction between the tire and the ground. It should be noted, of course, that there is a relation between the moment of inertia, the free tire radius and the aircraft weight in general the free tire radius and the moment of inertia will increase with aircraft weight. For any wheel and tire unit, it is shown that there are various combinations of landing speed and coefficient of friction which cause the wheel spinning up to just cease at the same instant as the maximum vertical wheel reaction is reached, and except for very gentle l]ridirs, the maximum value of 1 re quired is usually much less than 1. 0. N163404 Aeronautical Research Council (Gt. Brit.) THE SOLUTION BY LIF TINULINE THEORY OF PROBLEMS INVOLVING DISCONTINUITIES. V. M. Falkner. 1952. 23p. diagrs., 15 tabs. (ARC R 9 M 2592. Formerly ARC 10, 922; S & C 2158; Perf. 366) The report, which has been written as a preliminary to a later account of similar work in liftingplane the ory, describes how wing loading problems involving discontinuities are solved by liftingline theory. The four discontinuities considered are (a) direction of leading or trailing edge, (b) incidence, (c) two dimensional lift slope, and (d) chord. As the effects of the first are of minor importance in liftingline theory, attention is mainly confined to the last three, the solution being based on the use of a few terms of a Fourier series in conjunction with special functions tabulated elsewhere. The work is limited to straight unyawed flight and includes lift, induced drag, and pitching, rolling, and yawing moments, all with or without deflected landing flaps and ailerons. The method of formation of the equations, and the solu tions of a representative range of problems for a hypothetical wing, including loading due to incidence, symmetrical wing twist, uniform roll, and deflected flaps and ailerons, are fully described. An indica tion is given of how induced drag and yawingmoment calculations will later be simplified by the use of special derived functions. Absolute values of wing properties as given by liftingline theory are usually too high, but the specification of correction factors for viscosity is beyond the scope of the report. N16341* Aeronautical Research Council (Gt. Brit.) CALCULATED LOADINGS DUE TO INCIDENCE OF A NUMBER OF STRAIGHT AND SWEPTBACK WINGS. V. M. Falkner. WITH APPENDIX. Doris Lehrian. 1952. 43p. diagrs., 43 tabs. (ARC R S M 2596. Formerly ARC 11, 542; Perf. 449; S & C 2223) NACA RESEARCH ABSTRACTS NO.27 In this report are collected lriether the calculated aerodynamic loadings due to incidence of a number of straight and sweptback wings. The calculations follow in the main the routine described previously in another report, but include additions concerned with induced camber and induced drag. An additional in vestigation is made of the effect of the NACA camber on the properties at zero lift of a rectangular wing of aspect ratio 6. A table is given of loading functions for use in auxiliary solutions when the wing plan has a discontinuity of direction at an arbitrary position along the span. N16342 * Aeronautical Research Council (Gt. Brit.) ON THE DESIGN OF AEROFOILS FOR WHICH THE LIFT IS INDEPENDENT OF THE INCIDENCE. B. Thwaites. 1952. 17p. diagrs. (ARC R & M 2612. Formerly ARC 10, 294; FM 1057; Perf. 280) It has been shown in R & M 2611 how lift may be ob tained on airfoils independently of the incidence. In this paper mathematical processes are set out of de signing such airfoils to have specified velocity dis tributions at certain incidences and lift coefficients. Approximate and exact methods are given, corre sponding to the methods employed in the design of ordinary airfoils. Several shapes are worked out, some of them being the product of ideas not given in R & M 2611. A full discussion of the characteristics of such airfoils is given. N 16343* Aeronautical Research Council (Gt. Brit.) SPEEDS AND NORMAL ACCELERATIONS OF BOE ING CLIPPER AIRCRAFT ON NORTH AND SOUTH ATLANTIC ROUTES. D. T. Jones. 1952. 14p. diagrs., 6 tabs. (ARC R & M 2633; ARC 11, 640. Formerly RAE Structures 25) This report presents results obtained from Vg re corders fitted to Boeing Clipper aircraft on the North and South Atlantic routes between September 1944 and May 1946. The records cover about 3, 300 flying hours and show that the maximum speed recorded is 215 mph (I. A. S.) and the maximum upward and down ward accelerations are 2. 3g and 0. 3g, respectively. The two main groups of records considered differ from one another not only in respect of route but also in seasonal conditions and in proportion of flights made in wartime. Therefore, differences between the results cannot be simply ascribed to differences of route. It appears from the analysis that the maxi mum speed likely to be attained in a large flying time is somewhat greater in one group (North Atlantic) than in the other (South Atlantic) and that the maximum ac celerations on the other hand are likely to be less for the former than the latter. 9 N16353 Aeronautical Research Council (Gt. Brit.) THE MEASUREMENT OF YAWING MOMENT OF AN ELLIPTIC WING ON THE WHIRLING ARM. A. S. Halliday. August 2, 1949. 8p. diagrs. (ARC 12, 497; S & C 2320) The report describes tests on an elliptic wing on the Whirling Arm to determine the yawing moment due to a continuous rate of yaw. The results have been com pared with a theoretical estimation and the agreement found to be satisfactory. N163551 Aeronautical Research Council (Gt. Brit.) NOTE ON THE CONDITION OF MATRICES. Olga Todd. September 6, 1949. 3p. (ARC 12,561; 0.833) The object of this note is to establish the following theorem: let A be a real n x n nonsingular matrix and A' its transpose. Then AA' is more "ill conditioned" than A. This result confirms an opin ion expressed by Dr. L. Fox which he based on his practical experience. The term "condition of a ma trix" has been used in vague senses. The most com mon measure of a matrix has been the size of its de terminant: illconditioned matrices being those with small determinants. With this interpretation, the theorem is clearly correct. More adequate meas ures of the condition of a matrix have been proposed recently by John von Neumann and H. H. Goldstine and A. M. Turing. N16356* Aeronautical Research Council (Gt. Brit.) A NOTE ON THE DESIGN OF A HIGHSPEED WIND TUNNEL DRIVEN BY STEAM. T. R. CaveBrowne Cave. September 15, 1949. 5p. diagrs. (ARC 12,577; TP 292) A steamdriven supersonic wind tunnel has been de signed for, and is being built by, the University Col lege in Southampton. It has a 6 inch by 21/2 inch working section and will operate at a Mach number of 1. 4 for a period of 3 minutes. Since it takes about 5 minutes for the boiler to return to pressure, a 3 min ute test could be made every 8 minutes with this set up. Because the boiler is already available at the school, this design permits a substantial saving in time and money. Another advantage will be the com parative lack of noise as compared with a drive using a compressor. N16459* Royal Aircraft Establishment (Gt. Brit.) THE RHEOLOuY OF LUBRICANT FILMS. PART 2. E. Bielak and E. W. J. Mardles. February 1952. 27p. diagrs. (RAE Chem. 480) UNIVERSITY OF FLORIDA 3 1262 09079 7498 NACA 10 RESEARCH ABSTRACTS N0.27 This report presents further results on the rheolo gical properties of thin lubricant films (including greases) moving between surfaces, with special re card to the influence of the material of the surfaces on the rate of flow. The limitation of Amontons' law to pressures above 100 g/sq cm for lubricants is described in relation to the phenomenon of the "wring inf" together or "striction" of surfaces at light pres sures, especially with indifferent lubricants, when film rupture occurs early during the thinning of the film. The latent period, observed and first des cribed by Hardy, during which friction changes occur has been measured for several lubricants at low pres sures. Explanations of hydraulic "striction" and the "latent period" are suggested. MISCELLANEOUS NACA TN 2310 Errata No. I on "GENERALIZATION OF BOUNDARY LAYER MOMENTUMINTEGRAL EQUATIONS TO THREEDIMENSIONAL FLOWS INCLUDING THOSE OF ROTATINu SYSTEM". Artur Mager. March 1951. NA"AT.2naPleV 81552 4000 
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