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National Advisory Committee for Aeronautics Research Abstracts NO. 90 SEPTEMBER 27, 1955 CURRENT NACA REPORTS NACA Rept. 1197 A STUDY OF THE CHARACTERISTICS OF HUMAN PILOT CONTROL RESPONSE TO SIMULATED AIR CRAFT LATERAL MOTIONS. Donald C. Cheatham. 1954. 11, 14p. diagrs., photos., tab. (NACA Rept. 1197. Formerly RM L52C17) There are presented studies of the characteristics of pilot ability to control dynamically unstable yawing oscillations, studies of pilot control response to simulated aircraft yawing motions, and studies oi the feasibility of representing pilot control response in an analytical form. NACA RM E55F28a / STATISTICAL SURVEY OF 1 N fAURED ON SCHEDULED AIRLINE F 1HTS VER THE /  UNITED STATES AND CAN AIFOM NOVEMBEMt 1951 TO JUNE 1952. Porter ins. September 1955. 44p. diagrs., oto (NACA RM E55F28a)  A statistical survey and a preliminary analysis are made in an interim report of over 600 icing en counters obtained from a continuing program sponsored by the NACA with the cooperation of the airlines. Pressuretype icingrate meters were in stalled on 11 airline aircraft of various types. Icing conditions measured during scheduled operations gave relative frequencies of liquidwater content, icing rate, total ice accumulations, cloud tempera tures, as well as horizontal and vertical extent of icing clouds. Liquidwater contents were higher than data from earlier research flights in layertype coTdsbuttst~ghtty lower than previous data from cumulus clouds. NACA TM 1330 THEORY OF DYNAMIC CREEP. (K teorii dinamicheskoi polzuchesti). A. A. Predvoditelev and B. A. Smirnov. September 1955. 12p. diagr. (NACA TM 1330. Trans. from Moscow Universitet, Vestnik, v.8, no.8, 1953, p.7986) An analysis is given of the causes of the increase in creep under varying loads. It is suggested that the increase in creep is due to local rise in temperature over the slip planes, thus facilitating slip. A theory of dynamic creep is proposed based on theBecker  theory of the after flect, tdi8th treats the.rmetal as a granular structure and ittnCdes a rate factor. Comparison of the t eor th experitentat results is reserved for a fu ure apelr. ,., " L .. NACA TN 3293 CUMULATIVE FATIGUE DAMAGE OF AXIALLY LOADED ALCLAD 75ST6 AND ALCLAD 24ST3 ALUMINUMALLOY SHEET. Ira Smith, Darnley M. Howard, and Frank C. Smith, National Bureau of Standards. September 1955. 49p. diagrs., photos., 5 tabs. (NACA TN 3293) Results are presented of cumulativefatiguedamage tests made on 607 specimens machined from alclad 75ST6 aluminumalloy sheet 0.064 inch thick and 198 specimens of alclad 245T3 and alclad 755T6 aluminumalloy sheet 0.032 inch thick. The tests of the 0.064inchthick specimens 6lei'st ~5te 35 dif ferent loading conditions andln ,tl Ie.'4 32 inch material consisted ofI'S3 d rent loa ;n ditions. .SEP s 19551 / NACA TN 3294 ( 95 FRICTION STUDY OF AIRCRAFT"flREIAT IAL ON CONCRETE. W. G. Hamrite, Boeing Arfplane Company. September 1955. 34p. diagrs., photos. (NACA TN 3294) A systematic study was made of the variation of frictional resistance between typical tiretread material and three concrete surfaces of different roughness at various temperatures and normal pressures. The tiretread specimens were taken from the thickest portion of worn tenply tires, and the three concrete test blocks were poured from the same mix but subjected to dilerent surface finishes. Curves are presented ol the apparent coefficient of friction as a function of normal pressure. NACA TN 3477 HYDRODYNAMIC PRESSURE DISTRIBUTIONS OB TAINED DURING A PLANING INVESTIGATION OF FIVE RELATED PRISMATIC SURFACES. Walter J. Kapryan and George M. Boyd, Jr. September 1955. 82p. diagrs., photos., 5 tabs. (NACA TN 3477) *AVAILABLE ON LOAN ONLY ADDRESS REQUESTS FOR DOCUMENTS TO NACA, 1519 H ST., NW., WASHINGTON 25, D C., CITING CODE NUMBER ABOVE EACH TITLE. THE REPORT TITLE AND AUTHOR. :t? "/3 36 p' ceP w 2 Hydrodynamic pressure distributions have been ob "tained during pure planing for five related prismatic surfaces. The distributions gave integrated lifts that in almost every case were well within 10 percent of the applied load. Comparison of experiment with theory shows that existing theories will adequately predict flatplate pressures. For the Vshaped sur faces, experiment and theory are in poor agreement. The lift and centerofpressure data for both the flat and Vshaped surfaces are in good agreement with recent experimental and theoretical NACA research on planing surfaces. NACA TN 3479 ANALYSIS OF THE HORIZONTALTAIL LOADS MEASURED IN FLIGHT ON A MULTIENGINE JET BOMBER. William S. Aiken, Jr. and Bernard Wiener. September 1955. i, 69p. diagrs., photo., 6 tabs. (NACA TN 3479) Horizontaltail loads were measured in gradual and abrupt longitudinal maneuvers on two configurations of a fourengine jet bomber. The results obtained have been analyzed to determine the flight values of the coefficients important in calculations of hori zontal tail loads. The leastsquares procedure used to determine aerodynamic tail loads from strain gage measurements of structural tail loads which were affected by temperature is covered in detail. The effect of fuselage flexibility on the airplane motion is considered in the analysis of the abrupt maneuver data. When possible, windtunnel results are compared with flight results. Some calculations of critical horizontaltail loads beyond the range of the tests are given and compared with design loads. NACA TN 3486 MEASUREMENTS OF TURBULENT SKIN FRICTION ON A FLAT PLATE AT TRANSONIC SPEEDS. Raimo J(aakko) Hakkinen, California Institute of Technology. September 1955. 41p. diagrs., photo, tabs. (NACA TN 3486) The design and construction of a floatingelement skinfriction balance are described. This instru ment was applied to measurements of local skin friction in the turbulent boundary layer of a smooth flat plate at highsubsonic Mach numbers and super sonic Mach numbers up to 1. 75. The principal difficulties which exist in comparing skinfriction coefficients at various Mach numbers are discussed. NACA TN 3491 EXPERIMENTAL INVESTIGATION OF ECCENTRI CITY RATIO, FRICTION, AND OIL FLOW OF LONG AND SHORT JOURNAL BEARINGS WITH LOAD NUMBER CHARTS. G(eorge) B. DuBois, F(red) W. Ocvirk, and R. L. Wehe, Cornell University. September 1955. 63p. diagrs., tabs. (NACA TN 3491) NACA RESEARCH ABSTRACTS NO.90 The performance of plain bearings under steady central loading are compared and sunniarized by singleline curves covering the range of length diameter ratios both abore and below I. Experi  mental date on eccentricity ratio, friction, and oil flow for lengthdiameter ratios of 1, 11 2, and 2 are shown for comparison .'ith earlier data for lengthdiameter ratios of 1.4, 1 2. and 1. The combined data provide charts of plainoearing per formance which cover the practical range of length diameter ratio. NACA TN 3493 DEVELOPMENT OF EQUIPMENT AND OF EXPERI MENTAL TECHNIQUES FOR COLUMN CREEP TESTS. Sharad A. Patel, Martin Bloom, Burton Erickson, Alexander Chaick and N(icholas) J(ohn) Hoff, Polytechnic Institute of Brooklyn. September 1955. 20p. diagrs.. photos., tab. INACA TN 3493) Equipment and procedures developed for testing aluminumalloy columns subjected to constant loads at elevated temperatures are described. Particular emphasis was put on determination of the influence of initial dei nations from straightness on the critical time of the column, that is, the time necessary for the column to buckle %hen subjected to a constant load. Results are presented of tests of a number of 2024T4 aluminumnalloy columns having large slen derness ratios. NACA TN 3503 REDUCTION OF PROFILE DRAG AT SUPERSONIC VELOCITIES BY THE USE OF AIRFOIL SECTIONS HAVING A BLUNT TRAILING EDGE. Dean R. Chapman. September 1955. 29p. diagrs., photo. (NACA TN 3503. Supersedes RM A9H11) A preliminary theoretical and experimental investi gation has been made on the aerodynamic character istics of blunttrailingedge airfoils at supersonic velocities. The theoretical considerations indicate that properly designed airfoils with moderately blunt trailing edges can have less profile drag, greater liftcurve slope, and a higher maximum liftdrag ratio than conventional sections. These predictions have been substantiated by experimental measure ments on airfoils of 10percentthickness ratio at Mach numbers of 1.5 and 2.0, and at Reynolds num bers between 0. 2 and 1. 2 million. NACA TN 3514 RESPONSE OF HOMOGENEOUS AND TWO MATERIAL LAMINATED CYLINDERS TO SINUSOI DAL ENVIRONMENTAL TEMPERATURE CHANGE, WITH APPLICATIONS TO HOTWIRE ANEMOM ETRY AND THERMOCOUPLE PYROMETRY. Herman H. Lowell and Norman (A.) Patton. September 1955. ii, 143p. diagrs., tabs. (NACA TN 3514) NACA RESEARCH ABSTRACTS NO. 90 A theoretical investigating ot the response ot homno geneous and t.oniaterial laminated, iiiiiiite cylin ders to sinusoidal environnenrtal temperature and or small heattransfer coefficient changes was made. Generalized results are given ior the cylinder con sisting ol a shell of high thermal conductivity and a core of Io., conduct i. it.' The Othatnor ol a nuliber of specific platinurmfusedquartz ires" no '.arirng construction and diameter exposed to a representa tive airstream is indicated. For ratios l metal tnickness to overall radius of 0. 1, response ampli tude gains of aDout 4. 5 are predicted as compared with gains of more than 10 for infinitesimal shells. For a relative shell thickness of. 05.0 frequency re sponses at hot.ire aneiior..eters, exposedwire re sistance thermometers, or thermocouples Aould be extended by at least an order ol magnitude. Simnpli fied analyses are included which are not exact but are adequate lor design use. NACA TN 3522 MEASUREMENTS OF THE EFFECTS OF FINITE SPAN ON THE PRESSURE DISTRIBUTION OVER DOUBLEWEDGE WINGS AT MACH NUMBERS NEAR SHOCK ATTACHMENT. Walter G. Vincenti. September 1955. 50p. diagrs. (NACA TN 3522) Results are presented of measurements at low super sonic speeds of the pressure distribution on t1o wings having a common doublewedge section and aspect ratios 2 and 4. Comparable results for as pect ratio infinity have been published in NACA TN 3225. The results cover the Mach number range from 1.166 to 1.377, which brackets the value (1.221) for bowwave attachment at zero angle of attack. The data are discussed and compared with the previ ous twodimensional findings. NACA TN 3523 THE EFFECTIVENESS OF WING VORTEX GENERA TORS IN IMPROVING THE MANEUVERING CHARAC TERISTICS OF A SWEPTWING AIRPLANE AT TRANSONIC SPEEDS. Norman M. McFadden, George A. Rathert, Jr., and Richard S. Bray. September 1955. 43p. diagrs., photos., tab. (NACA TN 3523. Supersedes RM A51J18) The effects of wing vortex generators, multiple boundarylayer fences, and extension of the outer two segments of the wing leadingedge slats on the aerodynamic characteristics ot a 350 sweptwing fighter were measured in (light tests at transonic speeds and high altitudes. Significant improvements were obtained in the pitchup and wingdropping tendency characteristics with certain arrnagements of vortex generators. NACA TN 3562 VARIATION OF BOUNDARYLAYER TRANSITION WITH HEAT TRANSFER ON TWO BODIES OF REVOLUTION AT A MACH NUMBER OF 3.12. John R. Jack and N. S. Diaconis. September 1955. 16p. diagrs., photos. (NACA TN 3562) Cooling a conecylinder model to a walltofree stream ratio of approximately 1.4 increased the transition Reynolds number from a value of 2.0 x 106 at equilibrium to 10.6 x 106. For temperature ratios less than 1.4, the boundarylayer flow was en tirely laminar. For a parabolicnosed body, the transition Reynolds number was about twice that of the conecylinder model over the temperature range investigated. NACA TN 3563 HEAT LOSS FROM YAWED HOT WIRES AT SUB SONIC MACH NUMBERS. Virgil A. Sanaborn and James C. Laurence. September 1955. 44p. diagrs., photo. INACA TN 3563) Heatloss data at angles of yavw and fixed subsonic Mach numbers for several vires of different diam eters commonly used in hotmire anenomnetry are presented. Possible methods of correlating the data are examined. The relation oi the Reynolds number normal to the flo.., which has been used by most researchers, was inadequate except near a Mach number of zero. An empirical relation based on weighted addition of the heat losses of wires normal and parallel to the flow correlated all data reasonably well. NACA TN 3566 A POLARCOORDINATE SURVEY METHOD FOR DETERMINING JETENGINE COMBUSTION CHAMBER PERFORMANCE. Robert Friedman and Edward R. Carlson. September 1955. 29p. diagrs., photo., tab. (NACA TN 3566) An automatic polarcoordinate traversing system is described that sweeps a probe through a quarter annular exhaust duct circumferentially at selected radial positions. With a single combined pressure and temperature probe, temperature and pressure are recorded simultaneously as a function of probe position. The use of these data in calculating temperature and flow profiles, combustion efficiency, and pressure loss is shown. BRITISH REPORTS N38605* Aeronautical Research Council (Gt. Brit.) THE USE OF QUARTZ IN THE MANUFACTURE OF SMALL DIAMETER PITOT TUBES. J. R. Cooke. 1955. 14p. diagrs., photos., tab. (ARC CP 193) This note describes the method of manufacture of small quartztipped pitot tubes (down to 0.005 in. outside tip diameter) which have been used success fully for boundarylayer measurements on small models in a supersonic wind tunnel. Tests have been made of the effects of taper and end finish on the accuracy of measurement, and of the effect of the inside diameter of the tip (for a standard taper) on response rate. For a given inside tip diameter the tapered quartz tubes gave a faster response rate than the stainless steel hypodermic tubes previously used. N38606* Aeronautical Research Council (Gt. Brit.) A NOTE ON THE SOUND FROM WEAK DISTURB ANCES OF A NORMAL SHOCK WAVE. Alan PowelL 1955. 10p. diagrs. (ARC CP 194) The disturbances of a shock wave by sound waves or temperature fluctuations are studied in one dimen sion to a firstorder approximation. In general, both sound waves and temperature fluctuations arise behind the shock wave. Expressions are given for their amplitudes and calculated for y = 1.4. Sound waves colliding with the shock wave are amplified, but sound waves are almost annihilated by weak shock waves if originally travelling in the same di rection as the shock wave. Small temperature fluctuations give rise to much sound on an acoustical scale. N38607* Aeronautical Research Council (Gt. Brit.) REQUIREMENTS FOR UNIFORMITY OF FLOW IN SUPERSONIC WIND TUNNELS. D. E. Morris and K. G. Winter. 1955. 9p. diagr. (ARC CP 197) An analysis is made of the effects of nonuniformity of flow on the pressure measurements on the surface of a model and also on the force and moment meas urements. The following standards of flow uniform ity are derived variations in flow direction to be less than 0.10 in the range M = 1.4 to 3; variation in Mach number to be less than 0.003 at M = 1.4 increasing to 0.01 at M = 3. A brief analysis is made of the errors in model manufacture and their effects on force and pressure measurements. Using the same standards as were used in deducing the requirements for flow uniformity quoted above, it is concluded that present standards of model manu facture are satisfactory overall, though for accurate pressure plotting tests at low supersonic Mach num bers a higher standard is desirable. N38608* Aeronautical Research Council (Gt. Brit.) A CRITERION FOR THE PREDICTION OF THE RE COVERY CHARACTERISTICS OF SPINNING AIR CRAFT. T. H. Kerr. 1955. 22p. diagrs., tabs. (ARC CP 195) NACA RESEARCH ABSTRACTS NO. 90 It has been deduced that the t. o miost in.portant pa rameters are the unbalanced rollingnion.ent coeffi cient about the wind axis in the spin and the ratio of pitching to rolling moment of inertia. Using the results of fullscale spinning tests on 33 aircraft, it has been possible to establish empirical relation ships between the estimated unbalanced rolling moment coefficient and the inertia ratio which effect ively divide the aircraft into the three groups which have satisfactory, borderline, and unsatisfactory re covery characteristics. A simple method is pre sented for estimating the unbalanced rollingmoment coefficient knowing only the shape of the aircraft. The empirical relationships should give a good indi cation of the spinrecovery characteristics on new designs. N38616* Aeronautical Research Council (Gt. Brit.) MODEL TESTS ON THE EFFECTS OF SLIPSTREAM ON THE FLOW AT VARIOUS TAILPLANE POSI  TIONS ON A FOURENGINED AIRCRAFT. PART I. TESTS WITH CONTRAROTATING PROPELLERS. D. E. Hartley, A. Spence, and D. A. Kirby. PART II. TESTS WITH SINGLE ROTATING PRO PELLERS. D. A. Kirby. 1955. 37p. diagrs., tabs. (ARC R & M 2747; ARC 12, 355; ARC 14, 166. Supersedes RAE Aero 2322; RAE Aero 2322a) Systematic windtunnel tests have been made to in vestigate the effects of slipstream on the flow near the tail plane of a typical civil transport with four contrarotating propellers. Tallplane height has been varied for each of several wingbody arrange ments; only one tail plane and one propeller position have been used. This report presents the main re sults in the form of changes in mean downwash angle and velocity at the tail plane, as functions of tail plane position, lift coefficient, and propeller thrust. N38617* Aeronautical Research Council (Gt. Brit.) DETERMINATION OF THE STRESS DISTRIBUTION IN REINFORCED MONOCOQUE STRUCTURES. PART I. A THEORY OF FLATSIDED STRUC TURES. L.S.D. Morley. 1955. 23p. dagrs., photos. (ARC R &M 2879; ARC 14,814. Superse des RAE Structures 120) This paper is concerned with the estimation of the stress distribution in the neighborhood of a discon tinuity in reinforced monocoque flatsided struc tures. A theory is given based upon a shell model possessing uniformly distributed stringers but dis crete ribs, which can serve as a basis for the prac tical solution of a wide range of flatsided struc tures such as rectangular or polygonal fuselages and wing boxes. N38618* Aeronautical Research Council (Gr. Brit.) THE THEORETICAL WAVE DRAG OF SOME BODIES OF REVOLUTION. L. E. Fraenkel. 1955. 26p. diagrs., tab. (ARC R & M 2842; ARC 14, 334. Supersedes RAE Aero 2420) NACA RESEARCH ABSTRACTS NO. 90 This report investigates the wave drag of bodies of revolution with pointed or opennose forebodies and pointed or truncated afterbodies. The "quasi cylinder" and"slenderbody" theories are reviewed, a reversibility theorem is established, and the con cept of the interference effect of a forebody on an afterbody is introduced. The theories are applied to bodies whose profiles are either straight or para bolic arcs, formulas and curves being given for forebody and afterbody drag, and for the interfer ence drag. The results of the tAo theories are com pared and are seen to agree well in the region of geometries where both theories are applicable. N38619* Aeronautical Research Council (Gt. Brit.) AN EXPERIMENTAL INVESTIGATION OF STRESS DIFFUSION IN NONBUCKLING PLATES. L. H. Mitchell. 1955. 20p. diagrs., photos. (ARC R & M 2878. Supersedes ARC 14,934; Strut 1540) This report provides experimental results for com parison with theoretical analyses of stress diffusion problems. The structures considered consist of plane reinforced sheet which has been assumed not to buckle. Symmetrical loads are applied to the edge booms connected to the sheet by continuous no slip joints. Attention is concentrated on the stress distribution near the ends of the parallel strips of plate. An outline of the existing theoretical work which' is applicable to this type of problem is given. The stringersheet theory is compared with the photoelastic results. Some attention is also given to transverse end stiffeners which seem to have little effect on the shear stresses. N38620* Aeronautical Research Council (Gt. Brit.) THE BOUNDARY LAYER WITH DISTRIBUTED SUC TION. M. R. Head. 1955. 100p. diagrs., photos, tabs. (ARC R & M 2783. Supersedes ARC 13, 897; FM 1547; Perf. 771) Experiments performed in flight at Reynolds num bers in the region of 3 x 106 have clearly demon strated the stabilizing effect of small amounts of distributed suction on the laminar boundary layer. In the absence of a pressure gradient and in adverse gradients similar to those occurring on a normal airfoil, transition of the boundary layer to the tur bulent form has been prevented by the use of such suction quantities as may be expected to lead to very considerable reductions in effective drag. It ap pears, however, that for extensive laminar flow to be achieved in this way, the surface must be free from such excrescences as would cause transition in the absence of suction. N38621* Aeronautical Research Council (Gt. Brit.) METHODS FOR CALCULATING THE LIFT DISTRI BUTION OF WINGS (SUBSONIC LIFTINGSURFACE THEORY). H. Multhopp. 1955. 96p. diagrs.. tabs. (ARC R & M 2884; ARC 13,439. Supersedes RAE Aero 2353) These methods for calculating the load distribution on wings of any plan forn are based on the concep tions of liftingsurface theory. Computer work time is shortened by careful choice of the positions of pivotal points, by plotting once for all those parts of the doAnwash integral which occur frequently and by a consequent application of approximate integra tion methods similar to those devised by the author for liftingline problems. The basis of the method is to calculate the local lilt and pitching moment at a number of chordwise sections from a set of linear equations satisfying the donnvash conditions at two pivotal points in each section. N3871 Aeronautical Research Council (Gt. Brit.) SIMPLE EVALUATION OF THE THEORETICAL LIFT SLOPE AND AERODYNAMIC CENTRE OF SYMMETRICAL AEROFOILS. H. C. Garner. 1955. 20p. tabs. (ARC R & M 2847. Supersedes ARC 14,337; Perf.847; S & C 2561) This paper presents a simple method of calculating theoretical values of the lift slope (al)T and the position of aerodynamic center hT in two dimensional incompressible flow. Starting with the ordinates of an airfoil, the method in section 3 pro vides first and second approximations to both de rivatives, which are compared with exact theory and other calculated values in Tables 2 and 3 for various symmetrical airfoils listed in Table I. In section 5 a correction to the first approximation is introduced so as to permit the evaluation of (al)T within 1,2 percent and hT within about 0.001 in less than a quarter of an hour. A complete illus trative calculation is set out in Table 4. N38712' Aeronautical Research Council (Gt. Brit.) BOUNDARYLAYER CONTROL FOR HIGH LIFT BY SUCTION AT THE LEADINGEDGE OF A 40 DEG SWEPTBACK WING. E. D. Poppleton. 1955. 38p. diagrs., tabs. (ARC R & M 2897; ARC 14,771. Supersedes RAE Aero 2440) Windtunnel tests on the 10percentthick, constant chord, aspectratio4.6 wing are discussed. Boundarylayer control was applied along the whole leading edge; a comparison was made between the effects of distributed suction and suction through a slot. A 45percent Fowler flap was used in some tests. The overall effect of the two systems was similar, giving an increase in CLmax by increas ing the stalling angle of attack and making the wing statically stable up to the stall, when there was a severe loss of lift. The tests were designed to de termine whether leadingedge suction would produce comparable increases in CLmax on swept wings and, also whether tip stall could be prevented. N38713* Aeronautical Research Council (Gt. Brit.) ON THE APPLICATION OF OBLIQUE CO ORDINATES TO PROBLEMS OF PLANE ELAS TICITY AND SWEPTBACK WING STRUCTURES. W. S. Hemp. WITH AN APPENDIX. S. R. Lewis. 1955. 46p. diagrs., tabs. (ARC R & M 2754; ARC 12,981. Supersedes College of Aeronautics Rept. 31; College of Aeronautics Rept. 44) Methods are discussed by which designers can solve problems of stress distribution and deflection for the case of sweptback wing structures whose ribs lie parallel to the direction of flight. The mathe matical basis is developed and formulas are derived. The results are applied to a uniform, symmetrical, rectangular section sweptback box. Theories of stress distribution and deflections are obtained for the case of loading by normal forces and couples applied to the ends of the box. The main results are then generalized to cover the case of a more representative wing structure. Functions useful in the application of the theory are given in an appendix. N38714* Aeronautical Research Council (Gt. Brit.) LOWSPEED TUNNEL MODEL TESTS ON TAIL PLANE ROLLING MOMENTS IN SIDESLIP. A Spence, J. W. Leathers, and D. A. Kirby. 1955. 20p. diagrs., tabs. (ARC R& M 2941; ARC 14,701. Supersedes RAE Tech. Note Aero 2123) Measurements were made of the effect of sideslip on the rolling moment on a 41.50 sweptback tail plane mounted at three heights on the fin of a model of a single jet aircraft with a 400 sweptback wing. Inci dence and tailplane setting were varied, and the ef fects of rudder deflection were obtained with the tail plane at the top of the fin. Brief results on a delta aircraft model with a delta tail plane at the top of the fin are also included. Values of the rolling moment on the tail plane were obtained from meas urements of the bending moment on the starboard half of the tail plane about a hinge just outside the fin. NACA RESEARCH ABSTRACTS NO. 90 N38715* Aeronautical Research Counc II IGt. Brit.) TWODIMENSIONAL CONTROL CHARACTER ISTICS. L. W. Bryant, A. S. Halliday, and A. S. Batson. 1955. 47p. diagrA. IARC R M 2730. Supersedes ARC 13,039; S & C 2385: ARC 13.065; S & C 2386) Researches on the lift, pitching moments, and hinge moments of airfoils with plain flaps have been car ried out at the National Physical Laooratory at a Reynolds number of about 106. The results have been presented in a generalized form, *hich shows promise of being applicable over a wide field. It appears that a suggestion due to Preston that the ratio of experimental lift slope IdCL dy = al) to the theoretical value (al)T, corresponding to the Joukowsky condition of flo* past the trailing edge, provides a criterion giving the combined effects of Reynolds number, transition points, and airfoil shape on dCL/dct, and is a ivry useful starting point for the estimation of control characteristics. N38716* Aeronautical Research Council IGt. Brit.) PERMISSIBLE DESIGN VALUES AND VARIABILITY TEST FACTORS. R. J. Atkinson. 1955. 20p. diagrs., tabs. (ARC R & M 2877; ARC 11,619; ARC 13,748. Supersedes RAE Tech. Note Structures 15; RAE Tech. Note Structures 61) For the design of structural elements it is postulated that: not more than 10 percent of any given design should have strength below the design value, and not more than 0.1 percent should have strength below 90 percent of the design value. This rule forms a working basis for the interpretation of tests on sta tistical lines. On the basis of a fixed probability the report deduces: expressions for the derivation of permissible design values from a given number of test results, the number of test results required so that the estimates of permissible design values can be regarded as sufficiently accurate, and the factor which should be applied to the results of tests on any number of similar components designed to meet a specified requirement. N38717* Aeronautical Research Council iGd. Brit.) IMPROVEMENTS IN THE FATIGUE STRENGTH OF JOINTS BY THE USE OF INTERFERENCE FITS. W. A. P. Fisher and W. J. Winworth. 1955. 17p. diagrs., photos., tabs. IARC R & M 2874: ARC 15,014. Supersedes RAE Structures 127) NACA RESEARCH ABSTRACTS NO. 90 Fatigue test results are given for aluminum alloy flat bars with a single hole loaded by a pin in double shear. In one series the pin was fitted directly in S the hole with various degrees of interference fit up to 0.003 in. excess diameter. The other series had a mild steel bush interposed sith similar degrees of interference in the bar, but with a push fit between pin and bush. Both sets showed a great increase in fatigue strength for interference fits above a critical value. N38718* Aeronautical Research Council (Gt. Brit.) AN EXAMINATION OF THE FLOW AND PRESSURE LOSSES IN BLADE ROWS OF AXIALFLOW TUR BINES. D. G. Ainley and G. C. R. Mathieson. 1955. 33p. diagrs. (ARC R & M 2891; ARC 14,232. Supersedes NGTE R.86) Available information is studied and analyzed to de termine magnitudes of gas pressure losses and de Ilections in a wide variety of blade rows and to de termine the separate influences of variables such as blade shape, blade spacing, gas Mach number, Reynolds number, incidence, etc. Special attention is paid to "secondary losses. Effects of blade tip clearance are also considered. Empirical guiding rules and charts are derived from which approximate values of the overall pressure losses and gas deflec tions in a range of blade rows can be deduced. It is found that secondary losses can in many instances be large, the loss being generally found to be great when the blading has low reaction. N38719* Aeronautical Research Council (Gt. Brit.) FLUTTER AND RESONANCE CHARACTERISTICS OF A MODEL CANTILEVER WING CARRYING LOCALISED MASSES. N. C. Lambourne. 1955. 25p. diagrs., tabs. (ARC R & M 2866. Supersedes: ARC 13,910; 0.939; ARC 11,008; 0.687) Resonance tests on a model cantilever wing carrying concentrated masses were made in conjunction with flutter tests. Measurements were made with masses up to approximately five times the mass of the bare wing added at two positions. Flutter and resonance characteristics are placed in juxtaposi tion. An attempt is made to correlate the two sets of phenomena by means of the Kiissner criterion. Distortion modes of flutter are analyzed into normal mode components. Results suggest that for a wing rigidly fixed at the root and carrying a single con centrated mass the first three normal modes are sufficient to define the flutter mode. Copies obtainable from NACA, Washington 7 N38720' Aeronautical Research Council (Gt. Brit.) SOME APPLICATIONS OF THE LAME FUNCTION SOLUTIONS OF THE LINEARISED SUPERSONIC FLOW EQUATIONS. PART I FINITE SWEPT BACK WINGS WITH SYMMETRICAL SECTIONS AND ROUNDED LEADING EDGES. PART I CAMBER ED AND TWISTED WINGS. G. M. Roper. 1955. 42p. diagrs. (ARC R & M 2865; ARC 14,473; ARC 14,475; ARC 14,476. Supersedes RAE Aero 2436; RAE Aero 2437) In the present paper some special solutions are found. Some of these solutions are combined with previous solutions to give (a) pressure distribution and wave drag at zero lift on some finite unyawed sweptlack ings having symmetrical sections with rounded leading edges and wing tips perpendicular to the wind direction, and (b) the change in pressure distribution and wave drag at zero lift on the surface of a Sqwure wing when the thickness chord ratio is modiiied. Some additional solutions applicable to cambered and twisted wings are also given. N38721' Aeronautical Research Council (Gt. Brit.) THE APPLICATION OF THE EXACT METHOD OF AEROFOIL DESIGN. M. B. Glauert. 1955. 45p. diagrs., tabs. (ARC R & M 2683. Supersedes ARC 10,933; FM 1161) This report considers in detail the design of air foils by Lighthill's exact method, in which the ve locity over the airfoil surface is prescribed as a function of the angular coordinate on the circle into which the airfoil may be transformed. The mathe matical basis of the method is set out, means for obtaining desired characteristics for the airfoil are developed, and the procedure to be followed in the actual design is fully discussed. Various special functions are introduced to increase the range and practical utility of the velocity distributions obtain able, and these and other functions are fully tabu lated. The calculations for the design of a particu lar thick suction airfoil are set out in detail. Copies obtainable from NACA, Washington N38722* Aeronautical Research Council (Gt. Brit.) AN EXPERIMENTAL INVESTIGATION OF THE BOUNDARY LAYER ON A POROUS CIRCULAR CYLINDER. D. G. Hurley and B(rian) Thwaites. 1955. 14p. diagrs., photos. (ARC R & M 2829. Supersedes ARC 14,158; FM 1584) The report describes an experimental investigation of the boundary layer on the surface of a porous circular cylinder at which there is a normal inward velocity. The primary object of the experiments was to test the approximate theory of reference 1 for calculating the development of a laminar boundary layer under conditions of continuous suction. The formula given in that reference for calculating the momentum thickness of the layer gave results in ac cord with the experimental determinations. Owing to practical difficulties in the exploration of the very thin boundary layers and in the determination of the velocity gradient around the surface, other com parisons with the theory were difficult. N38723* Aeronautical Research Council (Gt. Brit.) FORMULAE FOR ESTIMATING THE FORCES IN SEAPLANEWATER IMPACTS WITHOUT ROTA TION OR CHINE IMMERSION. R. J. Monaghan and P. R. Crewe. 1955. 28p. diagrs., tabs. (ARC R & M 2804; ARC 12,399. Supersedes RAE Aero 2308) This report contains design formulas for estimating the maximum forces, together with the times and drafts associated with these forces, in mainstep landings of seaplanes provided there is neither ro tation nor chine immersion. Good agreement is formed with the results of model tests made under controlled conditions at NACA. The basic formulas and curves presented are considered to be the most satisfactory and accurate of the many proposed in recent years. They involve the use of a new basic parameter which is a measure of the effect of forward velocity; a new formula for associated mass, and a new method of plotting which is con sidered to be the most useful for the analysis of ex perimental data. N38724* Aeronautical Research Council (Gt. Brit.) WINDTUNNEL TESTS ON THE NACA 63A009 AEROFOIL WITH DISTRIBUTED SUCTION OVER THE NOSE. N. Gregory and W. S. Walker. 1955. 17p. diagrs., tabs. (ARC R & M 2900. Supersedes ARC 15,184; Perf. 987; FM 1787) The effects of distributed suction on the stalling characteristics of the airfoil are described. The most economical extent of suction was from the lead ing edge for 2.75 percent chord round the upper surface. At a R = 1.15 x 106, a suctionquantity coefficient of 0.0034 increased CLmax from 0.86 to 1.50 by delaying the stall from a = 110 to a = 200. Scale effect on the flow was investigated at a = 14. The airfoil was also tested with a 20percent split flap at 600 deflection. Suction gave half the increase on the flapped airfoil that it gave on the plain airfoil. The airfoil was modified for further testing by reducing the chord and blunting the nose. NACA RESEARCH ABSTRACTS NO. 90 N38725* Aeronautical Research Council (Gl. Brit.) DETAILED OBSERVATIONS MADE AT HIGH IN CIDENCES AND AT HIGHSUBSONIC MACH NUM BERS ON GOLDSTEIN 1442/1547 AEROFOIL. H. H. Pearcey and M. E. Faber. 1954. 52p. diagrs., photos., tabs. (ARC R & M 2849. Super sedes ARC 13,531; FM 1498; Perf. 714) Surfacepressure distribution, shockwave photo graphs, and observations of boundarylayer separa tion have been made over a wide range of angle of attack. The observations enable the effects of compressibility on CLmax and on the nature of the stall to be studied in detail for the twodimensional case. The pitchingmoment coefficients, also, can be integrated from the pressure distributions. Cer tain features of the results are thought to be of fairly general interest and application. N38728* Royal Aircraft Establishment (Gt. Brit.) TECHNIQUES FOR THE MEASUREMENT OF THE AERODYNAMIC FORCES ON OSCILLATING AERO FOILS. W. G. Molyneux. June 1955. 30p. diagrs. (RAE Tech. Note Structures 161) The various techniques for oscillatory force meas urements are considered in relation to their applica tion to the measurement of the aerodynamic coeffi cients for a rectangular wing oscillating in modes of vertical translation and uniform pitch. It is shown that the eight relevant coefficients Lz, Lz, L,, Li, Mz, M., My and M& are obtainable by any of the techniques described. The survey is not ex haustive, but it provides a basis for comparison of the various techniques and should be of assistance to investigators in this field in indicating the particular technique most likely to meet their requirements. N38729* Royal Aircraft Establishment (GL. Brit.) THE EFFECT OF WATER ON THE POROSITY OF PARACHUTE FABRICS. J. E. Swallow. May 1955. 18p. diagrs., tabs. (RAE Tech. Note Chem. 1248) Air flow through parachute fabrics was found to be seriously affected by water. The porosity of the nylon, cotton, Fortisan and Terylene fabrics ex amined was decreased and became negligible for the closer weaves. This was mainly a surface tension effect, but swelling was a contributory factor for cellulosic fabrics. Mockleno weave nylon fabrics were least affected. NACA RESEARCH ABSTRACTS NO. 90 N387301 Royal Aircraft Establishment (Gt. Brit.) ON THE INTEGRAL EQUATIONS OF TWO DIMEN SIONAL SUBSONIC FLUTTER DERIVATIVE THEORY. D. E. Williams. June 1955. 39p. (RAE Structures 181) This note gives the result of an attempt to find an analytical solution of Possio's integral equation  the equation which connects the downwash and the pressure distribution on an airfoil oscillating in two dimensional subsonic conipressible flow. A method is given for solving this problem and for solving the corresponding problem in incompressible flow the solution of Birnbaum's integral equation. N38732* Royal Aircraft Establishment (Gt. Brit.) THE DETERMINATION OF FLUORINE IN ORGANIC COMPOUNDS CONTAINING FLUORINE AND PHOS PHORUS. T. R. F. W. Fennell. May 1955. lip. diagr., tabs. (RAE Tech. Note Chem. 1251) A published method for the determination of fluoride in the presence of phosphate ion has been found to yield erroneous results. The method has been modified to overcome this fault. N38759* Aeroplane and Armament Experimental Establish ment (Gt. Brit.) THE EFFECT OF THE GROUND ON A HELICOPTER ROTOR IN FORWARD FLIGHT. I. C. Cheeseman and W. E. Bennett. July 11, 1955. 13p. diagrs. (AAEE'Res/288) An approximate method of estimating the effect of the ground on the lift of a rotor at any forward speed is described. Flight tests on several different air craft show reasonable agreement with the theory. Curves are given showing the relation between thrust, height, speed, and power. The theory has been extended to include the effect of a variation in blade loading and shows that within the range that this parameter takes on present single rotor heli copters the effect is small. N38761 Royal Aircraft Establishment (Gt. Brit.) A UNIFIED THEORY OF PERFECTLY PLASTIC PLATES. E. H. Mansfield. May 1955. 53p. diagrs. (RAE Structures 170) A theory is developed for determining the collapse load and the collapse mechanism for perfectly plas tic plates under normal loading. A number of solu tions to simple problems is first presented and the theory is extended to deal with plates of arbitrary plan carrying a concentrated load, and to plates of rectangular or regular polygonal plan carrying a uniformly distributed load. N38781 Forest Products Research Lao. (Gt. Brit.) INVESTIGATIONS INTO GLUES AND GLUING. PROGRESS REPORT EIGHTYFIVE JUNE 1955. BEHAVIOR OF GLUED WOOD PRODUCTS IN LIGHT NAVAL CRAFT. PART I SYNOPTIC REPORT. FIFTH YEAR'S ANALYSIS. R. J. Newall and L. S. Donian. 6p. (Forest Products Research Lao. Supersedes corresponding part of Progress Report 711 This investigation consists in storing samples of ply wood and other glued wood products in selected loca tions for periods up to 10 years. At Intervals, samples are removed and systematically tested for deterioration of the glue lines, iungal attack, etc. Inspections have been made at sixmonthly intervals over the past 5 years and a summary of the observa tions is presented. N38782' Forest Products Research Lab. (Gt. Brit.) COMPOSITE WOOD SECTION. TRIALS OF TIMBERS FOR PLYWOOD MANUFACTURE. ANINGUERIA ANINGUERIS ALTISSIMA UGANDA. (NO RELIABLE WEIGHT FIGURES AVAILABLE BUT PROBABLY BETWEEN 35 AND 40 LB. PER CUBIC FOOT AT 15 PER CENT MOISTURE CONTENT). (PROGRESS REPORT TWENTYEIGHT). June 1955. 12p. taos. (Forest Products Research Lab.) N38783' Forest Products Research Lao. (Gt. Brit.) COMPOSITE WOOD SECTION. TRIALS OF TIMBERS FOR PLYWOOD MANUFACTURE. ABURA (NZINGU)MITRAGYNA STIPULOSA UGANDA. (36 POUNDS PER CUBIC FOOT AT 15 PER CENT MOISTURE CONTENT). (PROGRESS REPORT TWENTYSEVEN). June 1955. 14p. tabs. (Forest Products Research Lab.) N38784' Forest Products Research Lab. (Gt. Brit.) COMPOSITE WOOD SECTION. TRIALS OF TIMBERS FOR PLYWOOD MANUFACTURE. DAHOMA PIPTADENIA AFRICANA UGANDA. 147 POUNDS PER CUBIC FOOT AT 15 PER CENT CONTENT). MUCHENCHE PIPTADENIA BUCHANANII UGANDA. (35 POUNDS PER CUBIC FOOT AT 15 PER CENT MOISTURE CONTENT). (PROGRESS REPORT TWENTYSIX). June 1955. lip. tabs. (Forest Products Research Lab.) N38785* Forest Products Research Lab. (Gt. Brit.) MOISTURE RELATIONS OF COMPOSITE WOOD PRODUCTS. PROGRESS REPORT TWENTY SEVEN JUNE 1955. THE FURROWING OF VENEERED BLOCKBOARD. J. F. S. Carruthers. 9p. diagrs., tabs. (Forest Products Research Lab. Supersedes Progress Report 26, May, 1954) UNIVERSITY OF FLORIDA 3 1262 08153 278 9 The purpose of this investigation was to determine the cause of the furrowing which sometimes occurs on the surface of veneered blockboard after polishing. Three different core constructions were employed and an explanation of the furrowing is given for each type. N38807* Royal Aircraft Establishment (Gt. Brit.) VELOCITY CALCULATIONS BY CONFORMAL MAP PING FOR TWODIMENSIONAL AEROFOILS. D. A. Spence and N. A. Routledge. February 1955. 48p. diagrs., tabs. (RAE Aero 2539) A method is derived for computing the conformal transformation between the plane of an airfoil of arbitrary shape (symmetrical or cambered), and the plane of its velocity potential at zero lift (in which the airfoil contour becomes a slit), in order to per mit calculations of the velocity at points off the sur face. The integral equation which relates the con tours is derived by an application of Cauchy's theorem, and solved to the order of the square of thickness ratio. The solution is found by repre senting the ordinate distribution by a Fourier series. The rapid tailingoff of the Fourier coefficients for all smooth airfoil shapes then leads to high accuracy being achieved with a comparatively small amount of effort. The method is straightforward and has proved easy to use. N38808* Royal Aircraft Establishment (Gt. Brit.) THE CHARACTERISTIC FREQUENCIES OF SMALL OSCILLATIONS IN THE FLOW PAST BLUFF BODIES. D. A. Spence. May 1955. 23p. diagrs. (RAE Aero 2532) Summary: When a bluff body is placed in a steady stream it experiences buffeting, the periodicity of which can be explained in terms of interactions between external and boundary layer regions. It is shown that the frequency must satisfy a character istic equation in order for the oscillations induced in the boundary layer to be compatible with those in the outside stream. The equation is derived formally for Lighthill's step case and for that of the circular cylinder. The Karman vortices which are observed in the latter case appear to be a consequence of the oscillatory character of the circulation around the cylinder. DECLASSIFIED NACA REPORTS NACA RM A54F28 ON THE RANGE OF APPLICABILITY OF THE TRANSONIC AREA RULE. John R. Spreiter. August 1954. 21p. (NACA RM A54F28) (Declassi fied from Confidential, 9/7/55) NACA RESEARCH ABSTRACTS NO. 90 Some insight into the range of appbcability of the transonic area rule has been gained by comparison with the appropriate similarity rule of transonic flow theory and with experimental data for a large family of rectangular wings having NACA 63AXXX profiles. NACA RM A54J07 THEORETICAL PRESSURE DISTRIBUTIONS FOR SOME SLENDER WINGBODY COMBINATIONS AT ZERO LIFT. Paul F. Byrd. January 1955. 39p. diagrs. (NACA RM A54J07) (Declassified from Confidential, 9/7/55) Theoretical calculations are made of the pressure distributions for some slender, symmetrical wing body combinations in subsonic and supersonic flow. The combinations consist first of nonlifting, swept back wings mounted on a circular cylinder and second of such wings mounted on a body indented so that the local crosssectional area of the combina tion is constant. The results indicate that indenta tion straightens out the isobars along the wing and diminishes the maximum perturbation velocities. NACA RM L52HO8 A STUDY OF THE ZEROLIFT DRAGRISE CHAR ACTERISTICS OF WINGBODY COMBINATIONS NEAR THE SPEED OF SOUND. Richard T. Whitcomb. September 1952. 41p. diagrs., photos., 3 tabs. (NACA RM L52H08) (Declassified from Confidential, 7/26/55) Results are presented which indicate that near the speed of sound the zerolift drag rise of a thin low aspectratio wingbody combination is primarily de pendent on the axial distribution of the cross sectional areas normal to the airstream. Results of an investigation of applications of this concept to the reduction of the dragrise increments of repre sentative wingbody combinations are also presented. NACA RM L54A29a ON SLENDERBODY THEORY AT TRANSONIC SPEEDS. Keith C. Harder and E. B. Kunker. March 1954. 12p. (NACA RM L54A29a) (Declassified from Confidential, 9/7/55) The basic ideas of the slenderbody approximation have been applied to the nonlinear transonicflow equation for the velocity potential in order to obtain some of the essential features of slenderbody theory at I ransonic speeds. The results of the investigation are presented from a unified point of view which demonstrates the similarity of slenderbody solu tions in the various Mach number ranges. The tran sonic area rule and some conditions concerning its validity follow from the analysis. NACA Langley Field, Va. 
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