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onal Advisory Committee for Aeronautics Research Abstracts N0.26 JULY 23, 1952 CURRENT NACA REPORTS NACA TN 2711 THE AERODYNAMIC DESIGN OF HIGH MACH NUM BER NOZZLES UTILIZING AXISYMMETRIC FLOW WITH APPLICATION TO A NOZZLE OF SQUARE TEST SECTION. Ivan E. Beckwith, Herbert W. Ridyard and Nancy Cromer. June 1952. 30p. diagrs., 5 tabs. (NACA TN 2711) A method is given for the design of threedimensional nozzles utilizing axisymmetric flow to produce uni form flow in a test chamber of arbitrary cross sec tion. The method is applied to obtain the coordinates of a Mach number 10 nozzle for which a square test section is specified. Radial flow is used in a portion of the flow field to reduce the computation time. The remainder of the flow field is computed by the method of characteristics, but a simplified method is used near the axis. Tables which facilitate computation of the radial flow and the flow near the axis are included. Transition streamlines determined from the analytic expressions of Kuno Foelsch are compared with the streamlines obtained from the characteristic net of the Mach number 10 nozzle. NACA TN 2720 EFFECTS OF ASPECT RATIO ON AIR FLOW AT HIGH SUBSONIC MACH NUMBERS. W. F. Lindsey and Milton D. Humphreys. July 1952. 10p. photos., diagrs. (NACA TN 2720. Formerly NACA RM L8G23) Schlieren photographs were used in an investigation to determine the effects of changing the aspect ratio from infinity to 2 on the air flow past a wing at high sub sonic Mach numbers. The results indicated that the decreased effects of compressibility on drag coeffi cients for the finite wing are produced by a reduction in the compression shock and flow separation. NACA TN 2724 TRANSONIC SIMILARITY RULES FOR LIFTING WINGS. Keith C. Harder. June 1952. 26p. diagr. (NACA TN 2724) Transonic similarity rules for lifting wings at small angles of attack are derived. These rules make it possible to present each aerodynamic coefficient for a family of wings of varying aspect ratio and thick ness ratio in a single design chart for the transonic range. NACA TN 2726 ON THE APPLICATION OF TRANSONIC SIMILAR ITY RULES. John R. Spreiter. June 1952. 45p. diagrs. (NACA TN 2726) The transonic aerodynamic characteristics of wings of finite span are discussed from the point of view of a unified small perturbation theory for subsonic, transonic, and supersonic flows about thin wings. This approach avoids certain ambiguities which ap pear if one studies transonic flows by means of equa tions derived under the more restrictive assumption that the local velocities are everywhere close to sonic velocity. The relation between the two methods of analysis of transonic flow is examined, the simi larity rules and known solutions of transonic flow theory are reviewed, and the asymptotic behavior of the lift, drag, and pitchingmoment characteristics of wings of large and small aspect ratio is discussed. It is shown that certain methods of data presentation are advantageous for the effective display of these characteristics. NACA TN 2727 EXPERIMENTS IN EXTERNAL NOISE REDUCTION OF A SMALL PUSHERTYPE AMPHIBIAN AIR PLANE. John P. Roberts and Leo L. Beranek, Aeronautical Research Foundation. July 1952. 142p. diagrs., photos., 3 tabs. (NACA TN 2727) As part of a program to find practicable ways of re ducing the external noise level of light airplanes, noise measurements were made on a representative pushertype amphibian airplane. Tests were made with a standard airplane and with modified versions of the airplane equipped with a geared engine, exhaust muffler, and propellers with various numbers of blades. Soundlevel recordings were made of take offs; of overhead flights at 100 and 500foot altitude; and, for some configurations, of flights at 500foot altitude passing 3000 feet away. The tests also in cluded analyses of soundfrequency components with the airplane on the ground from a distance of 50 feet and at various positions around the airplane. The results are compared with measurements from a previous noise study of a tractortype airplane. NACA TN 2730 CHOKING OF A SUBSONIC INDUCTION TUNNEL BY THE FLOW FROM AN INDUCTION NOZZLE. W. F. Lindsey. July 1952. 20p. diagrs. (NACA TN 2730) *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 A decrease in Mach number with increase in induc tion jet pressure has been shown experimentally, for one type of induction tunnel, to be associated with choking of the induced flow. The analysis of the flow conditions provides a means of predicting the occur rence of induction choking so that it can be adequately considered in design. NACA TN 2732 THEORETICAL INVESTIGATION OF VELOCITY DIAGRAMS OF A SINGLESTAGE TURBINE FOR A TURBOJET ENGINE AT MAXIMUM THRUST PER SQUARE FOOT TURBINE FRONTAL AREA. (Revised). Leo Cohen. June 1952. 34p. diagrs., tab. (NACA TN 2732) The aerodynamic requirements of a singlestage tur bine for a turbojet engine at maximum thrust per square foot turbine frontal area are presented in the form of velocity diagrams. An impulse diagram with supersonic velocities relative to the leading and trail ing edges of the rotor is required for practically all blade speeds and turbineinlet temperatures investi gated. NACA TN 2733 METHOD FOR CALCULATION OF HEAT TRANSFER IN LAMINAR REGION OF AIR FLOW AROUND CYL INDERS OF ARBITRARY CROSS SECTION (INCLUD ING LARGE TEMPERATURE DIFFERENCES AND TRANSPIRATION COOLING). E. R. G. Eckert and John N. B. Livingood. June 1952. 71p. diagrs. (NACA TN 2733) A method which permits approximation of local heat transfer coefficients in the laminarflow region around cylinders of arbitrary cross section from those for wedgetype profiles is extended to include the effects of large temperature differences and tran spiration cooling. Charts prepared from exact solu tions of the laminar boundarylayer equations for wedgetype profiles which allow for these effects yield results with a minimum of calculation. Application of the method to circular and elliptic transpiration cooled cylinders is made to determine local heat transfer coefficients and surface temperatures and to determine the variation in coolant flow required for maintaining a constant surface temperature. NACA TN 2735 AN ANALYSIS OF THE NORMAL ACCELERATIONS AND AIRSPEEDS OF A TWOENGINE TYPE OF TRANSPORT AIRPLANE IN COMMERCIAL OPERA TIONS ON ROUTES IN THE CENTRAL UNITED STATES FROM 1948 TO 1950. Walter G. Walker and Paul W. J. Schumacher. July 1952. 30p. diagrs., 4 tabs. (NACA TN 2735) Normalacceleration and airspeed data obtained from a twoengine type of transport airplane are analyzed to determine the gust and gustload experiences of the airplane. The gust and gustload experiences are compared with similar results from the same type of NACA RESEARCH ABSTRACTS NO.26 airplane during past operations. The influences of the gusts and the operating speeds in rough air on the gustload experience are indicated. NACA TN 2736 TWODIMENSIONAL SHEAR FLOW IN A 90 ELBOW. James J. Kramer and John D. Stanitz. July 1952. 44p. diagrs. (NACA TN 2736) As part of an approach to a better understanding of the motion of real fluids in flow machinery, two dimensional, incompressible, nonviscous shear flows in a 90 elbow have been investigated. Solutions are presented for linear and sinusoidal velocity distribu tions across the inlet of the elbow. The solutions with linear inlet velocity distributions indicate that as the negative vorticity of the flow increased: (1) the staticpressure drop through the elbow decreased, (2) the local deceleration along the outer channel walls increased, and (3) the magnitude of the veloci ties on the channel walls changed greatly, but the local pressure coefficient rose only gradually and the difference in pressure coefficient at corresponding points on the two walls was practically unchanged. In the case of a sinusoidal inlet velocity distribution, local decelerations occurred on both walls. NACA TN 2738 A PROBABILITY ANALYSIS OF THE METEORO LOGICAL FACTORS CONDUCTIVE TO AIRCRAFT ICING IN THE UNITED STATES. William Lewis and Norman R. Bergrun. June 1952. 93p. diagrs., 11 tabs. (NACA TN 2738) Meteorological icing data obtained in flight in the United States are analyzed statistically and methods are developed for the determination of (1) the various simultaneous combinations of the three basic icing parameters (liquidwater content, drop diameter, and temperature) which would have equal probability of being exceeded in flight in any random icing encoun ter, and (2) the probability of exceeding any specified group of values of liquidwater content associated simultaneously with temperature and dropdiameter values lying within specified ranges. NACA TN 2740 EXPERIMENTAL INVESTIGATION OF THE LOCAL AND AVERAGE SKIN FRICTION IN THE LAMINAR BOUNDARY LAYER ON A FLAT PLATE AT A MACH NUMBER OF 2.4. Randall C. Maydew and Constantine C. Pappas. July 1952. 22p. diagrs. (NACA TN 2740) Average and local skinfriction coefficients for lami nar flow have been determined experimentally on a flat plate at a Mach number of 2.4 for a Reynolds num ber range of 0.72 x 106 to 2.8 x 106 and compared with the laminarboundarylayer theory of Chapman and Rubesin. The average skinfriction data were considerably higher than values predicted by laminar boundarylayer theory. Local skinfriction crJeffi cients, evaluated by the two methods, agreed well with the theory when plotted against a momentum thickness Reynolds number. ^1 NACA RESEARCH ABSTRACTS NO.26 NACA TN 2742 BOUNDARYLAYER DEVELOPMENT AND SKIN FRICTION AT MACH NUMBER 3.05. Paul F. Brinich and Nick S. Diaconis. July 1952. 49p. diagrs., photos. (NACA TN 2742) Boundarylayer studies consisting of schlieren obser vations and momentum surveys were made on hollow cylinder models with their axes alined parallel to the stream. Results were obtained for three model diam eters and for natural and artificially induced turbulent boundarylayer flows. Transition Reynolds numbers were found to decrease with decreases in leading edge thickness and with reductions in tunnel pressure level. Turbulent temperaturerecovery factors gen erally decreased with increasing Reynolds number and were a maximum for the smallest transition Reynolds numbers. The results of this investigation appeared to be consistent with the theoretical turbu lent friction formulas of Wilson and with the extended FranklVoishel analysis of Rubesin, Maydew, and Varga. Velocity profiles in the outer portion of the boundary layer could be approximated reasonably with a 1/7 power profile and were found to be approxi mately similar in this region. Velocity profiles given by the Krmin universal turbulent boundarylayer Profile parameters were found to be similar in the laminar sublayer and in the turbulent region. NACA RM E52D17 EXPERIMENTAL INVESTIGATION OF AVERAGE HEATTRANSFER AND FRICTION COEFFICIENTS FOR AIR FLOWING IN CIRCULAR TUBES HAVING SQUARETHREADTYPE ROUGHNESS. Eldon W. Sams. June 1952. 43p. diagrs., photos. (NACA RM E52D17) A heattransfer investigation was conducted with air flowing through electrically heated Inconel tubes having various degrees of squarethreadtype rough ness (conventional roughness ratios of 0 (smooth tube), 0.016, 0.025, and 0.037), an inside diameter of 1/2 inch, and a length of 24 inches over ranges of bulk Reynolds number up to 350,000 and average inside tubewall temperatures up to 1950 R. The data showed that both heat transfer and friction increased with increase in surface roughness (particularly at the higher Reynolds numbers) and were also influ enced by tube walltobulk temperature ratio. Good correlation of the heattransfer data for all tubes was obtained by use of a modified Reynolds number includ ing the friction velocity. Correlation of the friction data for the rough tubes was obtained for the complete turbulence region by use of a modified correlation in corporating a roughness parameter which excludes the conventional roughness ratio; no overall corre lation was obtained for the region of incomplete turbulence. NACA RM E52D24 SOLUBILITY OF WATER IN HYDROCARBONS. R. R. Hibbard and R. L. Schalla. July 1952. 25p. diagrs., 3 tabs. (NACA RM E52D24) 3 An equation is presented which may be used to esti mate the solubility of water in any nonolefinic hydro carbon or hydrocarbon blend at any temperature. Only the temperature and the hydrogentocarbon ratio of the hydrocarbon need be known to calculate solubilities which are believed to be sufficiently accurate for most engineering purposes. NACA RM E52E19 USE OF FENCES TO INCREASE UNIFORMITY OF BOUNDARY LAYER ON SIDE WALLS OF SUPER SONIC WIND TUNNELS. Rudolph C. Haefeli. July 1952. 15p. diagrs., photos., tab. (NACA RM E52E19) An investigation of the use of solid fences installed on the side walls of a supersonic wind tunnel to re tard the development of transverse flow and thus to increase the uniformity of the sidewall boundary layer is reported. Beneficial results were obtained with fences which had depths of the order of the boundarylayer displacement thickness and which followed potentialflow streamlines through the nozzle. Reduction of the number of fences on each side wall from four to two eliminated their effective ness. NACA TM 1223 SOME EXPERIENCES REGARDING THE NONLINE ARITY OF HOT WIRES. (Quelques experiences sur la nonline'arit6 des fils chauds). R. Betchov and W. Welling. June 1952. 13p. diagrs. (NACA TM 1223. Trans. from Koninklijke Nederlandsche Akademie van Wetenschappen te Amsterdam, Ver handelingen, v.53, no.4, 1950, p.432439; Technische Hoogenschool te Delft, Laboratorium voor Aero en Hydrodynamica. Mededeling 66). Hotwire experiments illustrate some differences be tween the resistance characteristics as given by King's equation and those for a finite wire length. The experiments support previous theoretical work which accounted for a heat loss although the wire ends and a dependence of the convective heat loss on a non linear function of the temperature difference between the wire and the air. NACA TM 1346 NONLINEAR THEORY OF A HOTWIRE ANEMOM ETER. (Theorie nonlinealre de l'anemometre a fil chaud). R. Betchov. July 1952. 23p. diagrs. (NACA TM 1346. Trans. from Koninklijke Neder landsche Akademie van Wetenschappen te Amsterdam, Verhandelingen, v. 52, no. 3, 1949, p. 195207; Technische Hoogenschool te Delft, Laboratorium voor Aero en Hydrodynamica. Mededeeling 61). A theoretical analysis is presented for the hotwire anemometer to determine the differences in resist ance characteristics as given by King's equation for an infinite wire length and those given by the addi tional considerations of (a) a finite length of wire with heat loss through its ends and (b) heat loss due to a nonlinear function of the temperature difference between the wire and the air. 4 BRITISH REPORTS N6183 Ministry of Supply (Gt. Brit.) THE MECHANICAL CHARACTERISTICS OF PNEUMATIC TYRES. R. Hadekel. March 1960. iii, 146p. diagrs., photos. (MOS S The object of this report is to give an exposition of the existing kiiwlJ.dge of the mechanical properties. characteristics and behavior of pneumatic tires, in sofar as they may be of interest to the user. Ques tions of tire terhn,,h'y are not discussed. The roll ing ir_.,:ers. vertical and horizontal loaddeflection characteristics, behavior during braking and side ways motion, rolling resistance and high speed ef fects are all dealt with. Where possible, both theoretical treatment and empirical data are given. Recommendations are made with respect to further research. The primary object of the investigation was to provide data for aircraft purposes, but it is hoped that the information here presented will also be of use to land vehicle designers. N14407" Aeronautical Research Council (Gt. Brit.) SANDWICH CONSTRUCTION AND CORE MATERIALS. PART V. W. J. Pullen. 1952. 14p. diagrs., photos., 5 tabs. (ARC R : M 2686; ARC 10,501. Formerly ARC 9017; Strut 938; Plas. 57) Tensile, compressive and creep tests have been carried out on four different samples of extruded cellular cellulose acetate. It is concluded that the material is comparable with calcium alginate and other lowdensity materials so far handled in the Engineering Division, N.P.L. In particular, the samples are not subject to the same degree of "softening" as has been the case with some similar materials. The "filled" samples are more efficient than the "unfilled" ones and are worth considering as possible lowdensity stabilizers in sandwich construc tion. N14772* Aeronautical Research Council (Gt. Brit.) A REVIEW OF SOME STALLING RESEARCH. A. D. Young. WITH AN APPENDIX ON WING SECTIONS AND THEIR STALLING CHARACTERISTICS. H. B. Squire and A. D. Young. 1951. 39p. diagrs., 5 tabs. (ARC R & M 2609; ARC 5751. Formerly RAE Aero 1718) Over a period of years a considerable amount of stall ing research on various airplanes was completed at the Royal Aircraft Establishment and it was consid ered desirable that the main results should be sum marized and reviewed. The report includes a general discussion of the effect on stalling behavior of wing section, plan form, washout, flaps, nacelles, gills, slipstream, automatic wingtip slots and Hudson NACA RESEARCH ABSTRACTS NO.26 type slits. The important part that is played by the longitudinal trim and stability at incidences near the stall is emphasized. The relation between wing sec tions and their stalling characteristics is discussed and it is shown that the stalling characteristics can be broadly predicted from an examination of the form of the wingsection uppersurface pressure distribution at high incidences. The results indicate that vicious stalling behavior can be avoided by the use of wing sections towards the tip of fairly high camber (3 to 4 percent) and moderate thickness (> 12 percent). For some types of airplanes there are, however, serious objections to the use of high camber towards the tips; the designer is then advised to avoid wing sections which experiments and theory indicate have particularly bad stalling characteristics. The worst tip thickness for stalling appears to be in the region of 9 percent. High taper tends to worsen the stalling behavior and it is advisable to consider taper ratios greater than 2:1 only in conjunction with wingtip sections having good stalling characteristics. The use of partspan flaps does not appear to cause any marked deterioration in stalling behavior, and fre quently it improves the behavior; but there is some evidence, though not yet conclusive, that the use of fullspan flaps may be accompanied by an appreciable worsening in stalling behavior. Attention is drawn to the advisability of examining the flow at high inci dences in the neighborhood of the tailplane of an air plane in the design stage, with a view to assessing its probable stalling behavior; In particular, the possi bilities of designing for some stall warning can then be examined. N16033' Aeronautical Research Council (Gt. Brit.) THE DIFFUSION OF TRANSVERSE LOADS IN A REINFORCED CIRCULAR CYLINDER WITH NON RIGID FRAMES. S. R. Lewis. 1952. 23p. diagrs. (ARC CP 74) Explores the variation of shear stress distribution at a loaded frame consequent upon the variation of sev eral geometric parameters such as frame stiffness, skin thickness, stringer spacing, etc. Figures indi cate the shear distribution around a frame for a sin gle concentrated radial load of 1000 lb. The param eters chosen are those common to aircraft design, and it is possible to obtain a reasonably accurate shear distribution around a frame from the data sup plied, without doing the actual lengthy shear calcula tion. The work is confined to the particular case where the loaded frame is removed from the end of the cylinder. Gives method of obtaining the shear load due to a tangential load and moment from the radial load expressions. N16034* Aeronautical Research Council (Gt. Brit.) DIGITAL RECORDING AND ANALYSING OF FLIGHT TEST DATA: A PROPOSED SYSTEM. E. J. Petherick. 1952. 0lp. diagrs. (ARC CP 75) A small digital recorder is proposed, to punch 10,000 instrument readings on 100 feet of cine film, each NACA RESEARCH ABSTRACTS NO.26 item to 3 decimal or 12 binary places. Any number of instruments could be recorded at 10 readings a second. A further unit is envisaged, to read the punched data; to correct each item for instrument errors; to display the corrected values; and to punch them on Hollerith cards or RAESCC tape. N16036"* Aeronautical Research Council (Gt. Brit.) THE EFFECTS OF ATMOSPHERIC HUMIDITY AND TEMPERATURE ON THE ENGINE POWER AND TAKEOFF PERFORMANCE OF A HASTINGS I. G. Jackson. 1952. 34p. diagrs., 9 tabs. (ARC CP 77) Engine power decreases with increasing humidity and reduction is greatest at takeoff engine speed and boost. Specific humidity will rarely exceed 21/2 percent and this will decrease takeoff power approxi mately 10 percent as compared with dry air. The takeoff distance to clear a 50 ft. obstacle would be increased about 17 percent. For takeoff conditions, humidity is a parameter of the same order of impor tance as temperature. At constant humidity the rate of decrease of power with increase of temperature is not very different from the value given by the stand ard formula below fullthrottle height. No effect of humidity on fuel consumption has been detected. Fuel consumption decreases with increase in temperature. N16038* Aeronautical Research Council (Gt. Brit.) AN EXPERIMENTAL INVESTIGATION OF THE PERFORMANCE OF A PILOT PLANT FOR DRYING AIR BY SOLID GRANULAR ADSORBENTS. P. J. Bateman. 1952. 44p. diagrs., 2 tabs. (ARC CP 79) An investigation was undertaken to measure the per formance on a pilot plant scale of uncooled beds of granular adsorbent when drying air to the low humid ities required for supersonic wind tunnel operation. Two size grades of activated alumina and one of silica gel were tested in a bed 21/2 by 21/2 feet in plan, and of various depths between 12 and 25 inches. The bed was worked throughout at about atmospheric pres sure and air was drawn from atmosphere without treatment for humidity or temperature. The results are presented, in the main, as graphs of mean bed concentration mass of water adsorbed \mass of activated adsorbent) against inlet humidity for two values of the outlet humidity, 0.0005 and 0.0002 lb water/lb air. It is shown that over the ranges of the variables encountered the mean bed concentration decreases with increasing inlet humidity, increases with increasing bed depth at a fixed contact time, and decreases with increasing grain size. Although a definitive comparison is not possible there does not appear to be much difference in the performance of silica gel and activated alumina of the same grain size. 5 N16039T Aeronautical Research Council (Gt. Brit.) CALCULATED PRESSURE DISTRIBUTIONS FOR THE R.A.E. 100104 AEROFOIL SECTIONS. R. C. Pankhurst and H. B. Squire. 1952. 20p. diagrs., 3 tabs. (ARC CP 80) Gives the positions of maximum thickness, leading edge radii, trailingedge angles, coordinates, shapes, pressure distributions, and velocity distributions for the RAE 100104 airfoil sections. N16042* Aeronautical Research Council (Gt. Brit.) WIND TUNNEL TESTS ON A 90 APEX DELTA WING OF VARIABLE ASPECT RATIO (SWEEPBACK 36.80). PART I. GENERAL STABILITY. J. G. Ross, R. Hills and R. C. Lock. PART E. MEASUREMENTS OF DOWNWASH AND EFFECT OF HIGH LIFT DE VICES. R. C. Lock, J. G. Ross and P. Meiklem. 1952. 67p. diagrs., photos., 26 tabs. (ARC CP 83) Part I. Longitudinal and lateral stability measure ments have been made in a low speed tunnel on a delta wing of 90 apex angle with three different taper ratios. The tests included measurements with ground, the effect of a body, and measurements of elevon power. CLmax was 0.86 for all taper ratios but was max reduced to a trimmed value of 0.65 with a static mar gin of 0.10a, due to the large loss of lift caused by the elevons. A tip stall starts on the wings at a 8 to 12 depending on the taper ratio; this has compara tively little effect on pitching moments but a large ef fect on both rolling and yawing moments, nv and zv both decreasing after the tip stall. C. A. T. tests suggest that there is an appreciable favorable scale effect on the tip stall. Ground effects are small and can be estimated sufficiently accurately using existing theoretical work on unswept wings. Part 1. Wind tunnel measurements of downwash were made on delta wings of aspect ratios 4, 3, and 2.3, using a tail of delta plan form in three vertical positions at two chordwise stations behind the wing. The tests also included the effect of the tail, and of split flaps and nose flaps, on the stability near the stall and on CLmax The tip nose flaps proved effective in delay "max f ff ing the tip stall, and gave some increase in CLmax max With split flaps untrimmed CLmax was 0.95 and 1.2 with the flap in the forward and rear position respec tively. There was no change of trim with the flaps in the forward position with tail off; with the tail an in termediate flap position should give zero trim change. The downwash was large at high lift coefficients, ow ing to the early tip stall, and this caused a loss of tail efficiency with a corresponding slight instability near the stall, which should not be serious. A method is given of calculating the downwash at small incidences behind a delta wing, and the results show good agree ment with the measured values, provided that the ex perimental liftcurve slope is used. 6 N16047* Aeronautical Research Council (Gt. Brit.) A THEORETICAL INVESTIGATION OF THE RE SPONSE OF A HIGHSPEED AEROPLANE TO THE APPLICATION OF AILERONS AND RUDDERS. K. Mitchell, A. W. Thorpe and E. M. Frayn. 1952. 92p. diagrs., 6 tabs. (ARC R & M 2294; ARC 8831; ARC 9639. Formerly RAE Aero 2040; RAE Aero 1952) The response of a fast moving airplane to a lateral gust, and to applied rolling and yawing moments, is examined by means of the differential analyzer, taking a range of values of the principal lateral sta bility parameters, and including sufficient ranges of the other stability and inertia parameters to make the conclusions of general validity for highspeed flight. The motion following a sharpedged sidegust is shown to be a markedly oscillatory character, with an unpleasantly short period, particularly in small airplanes. The shortness of the period is probably the worst feature. A general survey is made of the dependence of the motion upon the various parameters, the differential analyzer results being supplemented by the use of approximate formulas, which were de veloped with a view to this application. Particular attention is paid to the amplitudes of the motion in roll, yaw, and sideslip, and it is seen that it may be difficult to make the motion less unpleasant. The period may be lengthened by reducing nv, but the improvement that is possible in this way is limited. Damping can be improved by reducing dihedral, or by increasing body side area: the addition of a forward fin, ahead of the center of gravity, would therefore be doubly helpful, lengthening the period and improving the damping. In studying response to applied moments attention is chiefly concentrated upon response to ailerons, and the theoretical results are compared with a theoretical standard motion produced by a con stant rolling moment together with a yawing moment varied so as to suppress sideslip. Response at high speeds is shown to be insensitive to changes in Zv and nv within their normal ranges, and good response to pure rolling moment is assured for all lateral sta bility characteristics other than those associated with the combination of small fin with large dihedral: this combination is worst at high values of the lateral rel ative density. The effect of adverse yawing moment from the ailerons is detrimental, and becomes worse as the dihedral is increased or fin area decreased. N16048* Aeronautical Research Council (Gt. Brit.) LATERAL RESPONSE THEORY. K. Mitchell. APPENDICES. E. M. Frayn. 1952. 57p. 8 tabs. (ARC R & M 2297; ARC 7933; ARC 8278; ARC 8544. Formerly RAE Aero 1925; RAE Tech. Note Aero 1513; RAE Tech. Note Aero 1584) It is usual, when investigating the response of an air plane to applied forces and couples, to examine its behavior when disturbed by rolling and yawing mo NACA RESEARCH ABSTRACTS NO. 26 ments, instantaneously or linearly applied. The relevant curves can be readily obtained by means of a differential analyzer: their calculation by existing methods is extraordinarily tedious, and analysis of lateral response, on a basis of the curves alone, a matter of some difficulty. The need was therefore felt for some new approach to the subject which would offer an alternative method of assessing response. A method is now given, whereby the calculation of re sponse to any forces whatever is made to depend upon the prior determination of a number of constants, for which the name "modal response coefficients"' is pro posed. The idea of the method is introduced by solving first by the method of variation of constants; later the relevant formulas are more simply obtained by operational methods. It is shown that the Laplace transformation scores over Heaviside's methods here in leading at once to a result of considerable general ity and in a form more convenient for computation. Methods for calculating model response coefficients and response curves are examined in appendices, recommended methods being presented as routine computations to be carried out on printed computing sheets (given as tables 1 and 2). "Optimuminterval" tables of certain functions occurring in the calculation of response curves are also given (tables 3 to 8). As a method of calculating response curves, even with the tabular aids presented here, the method of modal response coefficients is probably slightly less laborious than the usual methods: this work should be handled mechanically by the differential analyzer. The calculation of the modal response coefficients themselves, however, is not tedious, and may usefully be undertaken to aid in the study of curves calculated by other means. It is hoped that, when a little experi ence has been gained, it may prove possible to dis cuss response in terms of the modal response coeffi cients alone. N16049* Aeronautical Research Council (Gt. Brit.) RELAXATION TIME EFFECTS IN GAS DYNAMICS. J. C. Gunn. 1952. 44p. diagrs., 12 tabs. (ARC R & M 2338. Formerly ARC 9536; FM 910; TP 145; TJR 89) In the formulation of the equations of gas dynamics it is usual to make certain simplifying assumptions con cerning the properties of the gas under consideration. Thus, the equations are greatly simplified if it can be assumed that the gas is a perfect one, satisfying the equation p/p = RT, and also that the specific heat, Cp, is a constant. A further assumption, generally made, is that the energy content of the gas is capable of instantaneous change following the external condi tions, so that the internal energy is always a uniquely defined function of the temperature. These assump tions are, of course, often justified in practice, but there is need for caution in applying them to unfamil iar gases, or, as we shall see, to cases where the temperature range of the process is very wide. The object of this note, in particular, is to study the third of the assumptions referred to above, and we shall in vestigate the effect on various dynamical processes of NACA RESEARCH ABSTRACTS NO.26 the known time lag in the pickingup, by the vibra tional degrees of freedom of a gas, of their equilib rium energy. Three main cases will be considered: (1) flow in a wind tunnel, (2) flow round an airfoil, without shock waves, and (3) flow through shock waves. The case of propagation of sound in a medium exhibiting heat capacity lag is also briefly discussed. The aim of the consideration is twofold: firstly, to demonstrate how small the effects of heat capacity lag are in most aerodynamic phenomena, and secondly to indicate some special cases in which the effects be come more important. The preliminary physical data concerning the vibrational energy of a gas, and its adjustment to equilibrium, have been fairly fully pre sented in papers by Kantrowitz, and, more particular ly by Bethe and Teller2. As these papers may not be readily accessible the opportunity is first taken to summarize this basic physical knowledge. Some of the more general consequences of the heatcapacity lag are also deduced. N16050* Aeronautical Research Council (Gt. Brit.) WINGS OF FINITE ASPECT RATIO AT SUPERSONIC VELOCITIES. D. R. Taunt and G. N. Ward. 1952. 12p. diagrs. (ARC R & M 2421; ARC 9401. Formerly Department of Scientific Research and Experiment, Admiralty SRE/Airflow/29) An error in the wellknown work by Schlichting is corrected, and it is shown that the integral equation derived can be solved analytically. The main result is that for a plane rectangular wing of aspect ratio A, the lift coefficient is less than that for an infinite plane wing by a factor 1 1 where M is the 2A 2 1 Mach number of the undisturbed stream. It is shown that the analytical solution can be extended to plane trapezoidal wings of any cuttingoff angle. N16051* Aeronautical Research Council (Gt. Brit.) THE VALIDITY OF FULLSCALE PROPELLER BRAKING TESTS IN THE R. A. E. 24FT WIND TUNNEL. A. B. Haines and R. J. Monaghan. 1952. 9p. diagrs., 2 tabs. (ARC R& M 2473; ARC 10,309. Formerly RAE Tech. Note Aero 1851) During tests at negative pitch air flow through the working section is considerably disturbed. To check the effect of tunnel interference, tests were made in both the 5foot and 24foot wind tunnels. Tests also provided information on Reynolds number effects and solidity. It was concluded that the effects of tunnel interference can be allowed for by a correction to J. Over the range of reliable measurements this addi tivc correction decreases with J, becoming negligible for J > 1.6. For the tests an increase in Reynolds number had no appreciable effect on negative thrust. Correcting for a 2:1 increase in solidity by treating thrust and power as proportional to solidity may over estimate negative thrust for a given power by 30 per cent. 7 N16052* Aeronautical Research Council (Gt. Brit.) SUPERSONIC DIFFUSERS. J. Lukasiewicz. 1952. 30p. diagrs., photos., 3 tabs. (ARC R& M 2501; ARC 10,110. Formerly RAE Gas 8) Various possible types of supersonic diffuser are considered theoretically and the available experimen tal evidence is reviewed. The most obvious type of an efficient supersonic diffuser is the reversed super sonic nozzle. A stability criterion for the reversed supersonic nozzle is developed, which shows that the contraction between the entry and the throat is so limited that if a normal shock at entry is to be avoided such a diffuser is of little practical use unless an artificial means of inducing and maintaining the less stable flow can be found. Test results of simple pitot entries show that, in the absence of the boundary layer, the onedimensional theory holds for normal shocks. Since their efficiency is high for Mach num ber up to about 1.5, diffusers with normal shock at entry are quite satisfactory in this velocity range. The mechanism of the shock wave and boundarylayer interaction is considered. When normal shocks occur in the presence of the boundary layer, as in annular pitot entries and diffusers for supersonic wind tun nels, their compression efficiency is appreciably lower than the theoretical one. For very high veloc ities, exceeding M = 2.5, the multishock diffuser is practicable and offers high theoretical efficiency. The efficiencies of this type of diffuser are examined and shown to exceed 90 percent up to M = 3.0 for de signs having three or four oblique shocks. Various geometrical arrangements are considered from the point of view of reduction of frontal area for a given mass flow, and improvements on the single focused wave system are obtained. The high performance of multishock diffusers is confirmed by Oswatitsch's tests at Gottingen, the results of which are here summarized. The operation of supersonic diffusers at other than the design Mach number is briefly con sidered. N16053* Aeronautical Research Council (Gt. Brit.) NOTE ON THE FRISE AILERON. H. H. B. M. Thomas and E. R. Crabbe. 1952. 16p. diagrs. (ARC R & M 2502; ARC 10,213. Formerly RAE Aero 2163) In spite of its well known disadvantages the Frise aileron is still used and can yield a good aileron con trol when the balance required is not very close. There appears to be, therefore, a need for some means of estimating the hingemoment characteris tics of this type of aileron. The present note, besides giving a general discussion of the characteristics of Frise ailerons, includes a method of estimating the hinge moment from the hinge moment of the unbal anced ailerons. Means of improving the hinge mo ment characteristics, avoiding oscillation, and in creasing the rolling power of Frise ailerons are also discussed. These include the use of differentially geared tabs and slats attached to the wing or the aileron. 8 N16057* Aeronautical Research Council (Gt. Brit.) THE PREVENTION OF BINARY FLUTTER BY ARTIFICIAL DAMPING. R. A. Frazer. 1951. 16p. diagrs., 2 tabs. (ARC R & M 2552. Formerly ARC 7473; ARC 0.397) Formulas are obtained which provide an estimate of the amount of artificial control needed to prevent binary flutter. Results are presented in terms of a "minimum damping multiplier" defined as the ratio of the least direct damping coefficient required for absolute flutter prevention to the "natural" direct aerodynamic damping coefficient of the control sur face concerned. Numerical results are obtained for five different types of aircraft. N16058* Aeronautical Research Council (Gt. Brit.) WINDTULNNEL TESTS ON THE SPOILINu EFFECTS OF ENGINE COOLING GILLS ON RADIAL AIR COOLED INSTALLATIONS ON A WINu J. Seddon and J. A. Kirk. 1952. 26p. diagrs., 6 tabs. (ARC R & M 2558; ARC 5659. Formerly RAE Aero 1724) Spoiling drag associated with fully open gills at high CL can be very large if the gill exit is nearer to the wing leading edge than about 10 percent of the local wing chord; but the effect diminshes rapidly as this distance is increased. The drag due to spoiling is reduced if the cooling air is kept away from the nacellewing junction by emitting it at specified re gions around the exit, preferably at the bottom where the lift is a minimum. Larger gill angles would be needed to satisfy maximumflow requirements in this way. The returnflow cooling system, with nose exit, shows no evidence of large spoiling drag at high cooling flow. The data obtained might be useful for estimating effects of other forms of discharge of low energy air in front of a wing leading edge. N16059* Aeronautical Research Council (Gt. Brit.) THE PRODUCTION OF LIFT INDEPENDENTLY OF INCIDENCE THE THWAITES FLAP, PARTS I AND II. B. Thwaites. 1952. 20p. diagrs. (ARC R & M 2611. Formerly ARC 10,100; FM 1020; Perf. 252; ARC 11,023; FM 1176, Perf. 378) In Part I of this paper, the possibility of obtaining lift on a body in a uniform stream independently of the incidence is discussed, and a practical method which obtains this effect if given. It is shown that a small thin "flap" which may be moved about a wellrounded trailing edge through which, for example, continuous suction is applied will produce circulation about the airfoil. A necessary feature of this method is the prevention of separation of flow by boundarylayer suction, which is also used to reduce substantially the width of the wake. The method uses principles quite different from those which have been proposed in the past for obtaining increased lift Qn airfoils. The practical applications of the device are briefly dis NACA RESEARCH ABSTRACTS NO.26 cussed, and some interesting consequences pointed out. It will, for instance, be possible to fly with an airfoil always at zero incidence. Again, the stall in which the flow separates from near the leading edge may be completely avoided, for as the circulation and lift increase, the incidence may be decreased so that severe adverse velocity gradients occur nowhere but near the trailing edge. In Part II of the paper, a report is given of a prelimi nary experiment which was set up to investigate whether the theoretical predictions made about the efficacy of the flap were largely confirmed. A wholly porous circular cylinder was fitted with the flap and measurements were made of the pressure distribution round the cylinder for various positions of the flap. These observations showed that for angular deflection of the flap of less than 20, about 85 percent of the theoretical value of CL was realized: a maximum CL of about 5.6 was obtained. These results are taken to show that the physical principles of Part I are sound and that the Thwaites flap does, in fact, enable lift to be generated independently of incidence. N16060* Aeronautical Research Council (Gt. Brit.) NOTE ON THE CHARACTERISTIC CURVE FOR AN AIRSCREW OR HELICOPTER. C. N. H. Lock. 1952. 3p. diagr. (ARC R & M 2673. Formerly ARC 10,636; ARC H46) Modifies the method of plotting the curves as given in R & M 1026 and R& M 1014. The new variables are the square roots of the variables used in these earlier reports. The change of variables has the following three advantages: (1) The three principal working states now correspond to three different quadrants, (2) the representation in the neighborhood of the xaxis and the yaxis is more definite since the curve has a finite slope at both these points, and (3) the formulas of the vortex theory take simple forms. N16061* Aeronautical Research Council (Gt. Brit.) TANK TESTS ON A HULL WITH THE MAIN STEP FAIRED IN PLANFORM AND ELEVATION. D. I. T. P. LlewelynDavies. 1952. 36p. diagrs., photos., 4 tabs. (ARC R & M 2708; ARC 8807. Formerly RAE Aero 2029) Tank tests were required to find out whether the water characteristics of a hull with a main step, faired in both plan form and elevation, were comparable with those of a hull with a conventional Vee or transverse step. Stability diagrams and spray and resistance characteristics were obtained over a large range of loadings (CA0 = 0.616 to CA = 1.440). The fully ,\O faired step offers more possibility of designing a lon gitudinally stable flying boat hull than does the conven tional transverse or Vee step, but a hull with such a step is 5 to 10 percent less efficient hydrodynamically except at high speed. In order to avoid running too fine at high speed, it is recommended that the center of gravity should not be more than 0.46b ahead of the NACA RESEARCH ABSTRACTS NO.26 apex of the step. The modification to the step plan form makes little difference to the main spray char acteristics, but increase in allupweight reduces wing, tailplaue and propeller clearances. The effect of increase in load on the porpoising stability char acteristics is to raise both limits, with a tendency for the upper limit to rise more rapidly, but less regular ly, than the lower limit. The freetotrim attitudes also rise with increase in allupweight. The planing efficiency of the hull increases with increase of load, especially at high speeds. There is evidence of a second resistance hump at high speeds and also of a critical variation of planing efficiency with attitude under similar conditions. N16313" Royal Aircraft Establishment (Gt. Brit.) CHARTS OF THE WAVE DRAG OF WINGS AT ZERO LIFT. T. Lawrence. January 1952. 18p. diagrs. (RAE Tech. Note Aero 2139) Theoretical calculations of the wave drag at super sonic speeds of nonlifting wings of double wedge and biconvex section are reviewed, and the best method of presenting the results considered. Using this method, a representative selection of the available numerical evaluations of the theory is presented. These should be of value for wing drag estimation purposes. MISCELLANEOUS NACA TN 2657 Errata No. 1 on "SOME EFFECTS OF FREQUENCY ON THE CONTRIBUTION OF A VERTICAL TAIL TO THE FREE AERODYNAMIC DAMPING OF A MODEL OSCILLATING IN YAW". John D. Bird, Lewis R. Fisher and Sadie M. Hubbard. April 1952. UNPUBLISHED PAPERS N8809* THE PHUGOID OSCILLATION OF THE AIRPLANE WITH CONSIDERATION OF UNSTEADY AIR FORCES. (Die Phygoidschwingung des Flugzeugs bei Berick sichtigung instationarer Luftkrafte). J. Weissinger. March 1952. 33p. diagrs., 3 tabs. (Trans. from Zentrale fur wissenschaftliches Berichtswesen der Luftfahrtforschung, Berlin. FB 1430; Deutsche Versuchsanstalt fur Luftfahrt E. V., Berlin. Institute fuir Aerodynamik, March 29, 1941). An iteration method for determination of the influence of the unsteady air forces on the slow longitudinal oscillation of the airplane (to be calculated according to the method of small oscillations) is described and the results of a few numerical calculations are given. 9 N9665* SPECTRAL ENERGY DISTRIBUTION IN RADIATION DUE TO ENCOUNTERING SHOCK WAVES. (Reparti tion spectrale energetique dans la Lumiere emise lors de la Recontre d'Ondes de Choc). A. Michel Levy, H. Muraour and E. Vassy. April 1952. 19p. diagrs., photos. (Trans. from Revue d'Optique Theorique et Instrumentale, 1941, p.149160). This report deals with the determination of the energy distribution in a continuous spectrum ob tained at the encounter with shock waves in the at mosphere of various gases: argon, helium, oxygen, etc. The reference source is the positive crater of a carbon arc. Various regions of the gas chamber were explored. The radiation is not that of a black body. The present status of the theory of continuous spectra offers no possibility for interpreting this phenomenon. N12762 * THE STRESSSTROKE DIAGRAM OF AIRPLANE SHOCK ABSORBER STRUTS. THIRD PARTIAL REPORT THE LANDING IMPACT OF OLEO PNEUMATIC SHOCK ABSORBERS. (Zur kenntnis der Kraftwegdiagramme von Flugzeugfederbeinen, 3. Teilbericht: Der Landestoss von Olluftfederbelnen). K. Schlaefke. April 1952. 15p. diagrs., 3 tabs. (Trans. from Aerodynamische Versuchsanstalt Gottingen E.V., Technische Berichte, v. 11, no. 5, 1944, p. 137141). The stressstroke diagram of an oleopneumatic leg under the first impact at landing can be obtained by superposing onto the diagram of the undamped com pressed air leg, the damping diagram according to the second approximate method described in the pre ceeding partial report. The maximum stroke, max imum stress, and efficiency to be expected for divers damping, lift coefficients, and heights of drop are computed and represented diagrammatically for prac tical use. Lastly, it is indicated how the newly intro duced damping factor for oleopneumatic legs can be determined by experiment. N12763* ON THE RECIPROCAL EFFECTS BETWEEN SHOCK ABSORBER STRUT AND TIRE AT LANDING IMPACT OF AIRPLANE UNDERCARRIAGES. (Zur Kenntnis der Wechselwirkungen zwischen Federbein und Reifen beim Landestoss von Flugzeugfahrewerken). K. Schlaefke. May 1952. 18p. diagrs., 3 tabs. (Trans. from Aerodynamische Versuchsanstalt Gottingen E. V., Technische Berichte, v. 10, no. 11, 1943, p. 363367). This report deals with the reciprocal effects between shockabsorber strut and tire due to impact at land ing. It was found that the tire reaches its maximum springing earlier than the strut and that at the first instant of landing impact the energy is taken up solely by the tire. The results are represented in diagrams which can be used for the conversion of experimental data to other combinations of tires and shock absorber struts. UNIVI"SITY OF FLORIDA II III IIIl 11 I1I11I11III1111l l I1I1 11111l III II IIIIIIII 1 NACA 3 1262 09079 7506 NACA 10 RESEARCH ABSTRACTS NO.26 N15440* Massachusetts Inst. of Tech. ASYMPTOTIC SOLUTIONS OF THE STABILITY EQUATIONS OF A COMPRESSIBLE FLUID. Cathleen S. Morawetz. July 1951. 47p. diagrs. (Massachusetts Inst. of Tech.) Two conclusions of practical interest are reached here. It is proved rigorously that the proper branch of the multiplevalued asymptotic solutions has been used by Lees and Lin in NACA TN 1115. Establishes the existence or absence of an inner viscous region and its nature. N15499* HOTWIRE ANEMOMETER UTILIZIN, DIRECT CURRENT AMPLIFIER. (Chokur Yu Zofukuki No Riyoseru NessenFusikukeij 8p. diagrs. (Trans. of Jap. rept., February 26, 1944, WrightPatterson Field microfilm, Jap/WTD/Re/336). Des% ribeb the design of a hotwire anemometer for measuring the instantaneous velocity fluctuations by means of a direct current amplifier and gives the properties of the instrument as determined experi mentally. DECLASSIFIED NACA REPORTS NACA RM L8102 INVESTIGATION OF TWO PITOTSTATIC TUBES AT SUPERSONIC SPEEDS. Lowell E. Hasel and Donald E. Coletti. November 19, 1948. 24p. diagrs. (NACA RM L8102) (Declassified from Confidential, 6/4/52) The results of tests at a Mach number of 1.94 of an ogivalnose cylindrical pitotstatic tube and similar tests at Mach numbers of 1.93 and 1.62 of a service pitotstatic tube to determine body static pressures and indicated Mach numbers are presented and dis cussed. The radial pressure distribution on the cy lindrical bodies is compared with that calculated by an approximate theory. NACALangley 72352 4000 
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