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Research Abstracts 0 / SEPTEMBER 16, 1952 CURRENT NACA REPORTS NACA Rept. 1028 EFFECT OF ASPECT RATIO ON THE AIR FORCES AND MOMENTS OF HARMONICALLY OSCILLATING THIN RECTANGULAR WINGS IN SUPERSONIC PO TENTIAL FLOW. Charles E. Watkins. 1951. 17p. diagrs. (NACA Rept. 1028. Formerly NACA TN 2064) The linearized velocity potential for a thin, flat, rec tangular wing undergoing sinusoidal torsional oscil lations simultaneously with sinusoidal vertical trans lations in a supersonic stream is derived in the form of an infinite series in terms of a frequency param eter. Simple closed expressions that include the re duced frequency to the third power are given for the velocity potential, components of total force and mo ment coefficients, and components of section force and moment coefficients. It is found that the com ponents of force and moment coefficients for small aspectratio wings may deviate considerably from those for an infiniteaspectratio wing. The method of solution can be utilized for other plan forms, that is plan forms for which steadystate solutions are known. NACA Rept. 1053 INVESTIGATION OF TURBULENT FLOW IN A TWO DIMENSIONAL CHANNEL. John Laufer, California Institute of Technology. 1951. 20p. diagrs., photos. (NACA Rept. 1053. Formerly TN 2123) A detailed exploration of the turbulent flow character istics in a twodimensional channel is presented. The measurements were made at three Reynolds numbers, 12,300, 30,800, and 61,600, based on the half width of the channel and the maximrhum mean velocity. A channel of 5inch width and 12:1 aspect ratio was used for the investigation. NACA TN 2748 ON TRANSONIC FLOW PAST A WAVESHAPED WALL. Carl Kaplan. August 1952. 43p. diagrs. (NACA TN 2748) The simplified nonlinear differential equation for transonic flow past a wavy wall is solved by the meth od of integration a for 'se'el lger4i'.ip teurefo the solution of tesilting recurrence rrinhlas is shown and illust i% oy a number of eAx.'wes A numerical test of ln".''etence is applied ti a key poAer series in k. Bmnlirn '1e S i larXyr paran,  eter. and leads to the cluf tnibn t 'iooth sym metrical potential flow pa  'y wall is no longer possible when the critical value of the stream Mach number is exceeded. NACA TN 2749 1 ANALYSIS OF FLOW IN A SUBSONIC MIXEDFLOW IMPELLER. ChungHua Wu, Curtis A. Brown and Eleanor L. Costilow. August 1952. 38p. diagrs. (NACA TN 2749) A method recently developed for determining the steady flow of a nonviscous compressible fluid along a relative stream surface extending from hub to cas ing between two adjacent blades in a turbomachine is applied to investigate the through flow of air in an experimental mixedflow impeller of high solidity. The shape of the stream surface is taken to be the same as that of the mean camber surface of the blade which consists of all radial elements. The principal equation governing the through flow is solved by the relaxation method with the use of fourthdegree differentiation formulas for unequally spaced grid points caused by the varying hub and casing wall radii. A detailed analysis is made of both incompressible and compressible flow through the impeller, and contour plots of the stream function, velocity com ponents, total enthalpy, static pressure, and Mach number are presented and discussed. NACA TN 2762 AERODYNAMIC CHARACTERISTICS OF THREE DEEPSTEP PLANINGTAIL FLYINGBOAT HULLS AND A TRANSVERSESTEP HULL WITH EXTENDED AFTERBODY. John M. Riebe and Rodger L. Naeseth. August 1952. diagrs., photos., 5 tabs. (NACA TN 2762. Formerly RM L8127) An investigation was made to determine the aerody namic characteristics of three deepstep (92 percent of beam) planingtail flyingboat hulls differing only in the amount of step fairing and of a transversestep hull with extended afterbody. Minimum drag coeffi cients, which include the interference effects of the support wing, were 0. 0066 for the transversestep hull about the same as for a conventional hull and *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. CLASSIFIED DOCUMENT I hs doument contains classified itnforrmation afecting the Nattonal Defense of the United StatI w, trn the Inlb, ato n :.tn cqa tt to .y r' I t' r o period rn th a m l ry arlo ova. s; a ;'*'. )a rameanng of the Espionage Act, USC 0:31 and 32. Its tran or orthe rvelattono its Co .tent *t any r yanwr apppro t n ! fa n r oand t e y etr rro t a :tre . t  t an Unauthortied person Is prohibited by law. I Uhite s.Ictat zens o' knof s loyalty .m sreUh h o f ntcesl ul tf. n..r et3 !te . N0.O29 National Advisory Committee for Aeronautics 2 0. 0057 or 14 percent less for the hull with a deep un faired step. The hulls with step fairing had up to 44 percent less drag. Lii,,iluIhi.il and lateral stability was about the same as for a conventional hull. NACA TN 2763 GUSTRESPONSE ANALYSIS OF AN AIRPLANE IN CLUDING 'A ING BEND[NLu FLEXIBILITY. John C. Houbolt and Eldon E. Kordes. August 1952. 48p. di.irs., 3 tabs. (NACATN 2763) An i.il.'is is made of the gust response (including bending mnioment) of an airplane ha% ii, the degrees of freedom of vertical motion and wing I.,lnd.li.i' ii \iili ity and basic parameters are established. A con venient numerical solution of the response equations, wellsuited for trend studies, is developed and used in an example. A method is indicated for determin ing the gust causing a known response and a pro cedure is given for determining the response of an airplane directly from the known response of another airplane. NACA TN 2766 SOME EFFECTS OF AMPLITUDE AND I ILQUENCY ON THE AERODYNAMIC DAMPING OF A MODEL OSCILLATING CONTINUOUSLY IN YAW. Lewis R. Fisher and Walter D. Wolhart. September 1952. 24p. diagrs., photo. (NACA TN 2766) A fuselageverticaltail combination was oscillated in yaw through a range of amplitudes from 1 2 to 4 and a low range of the reducedfrequency param eter. The phase angles between the tail force and the displacement were measured and converted to values of the damping in yaw which are compared with the damping predicted by the unsteadylift theo ries and with the experimental steadystate daniping value. These tests were conducted at a Reynolds number of 442, 000. NACA TN 2769 EXPERIMENTAL AND THEORETICAL DETERMI NATION OF THERMAL STRESSES IN A FLAT PLATE. Richard R. Heldenfels and William M. Roberts. August 1952. 35p. diagrs., photo. (NACA TN 2769) Thermal stresses induced in a flat, rectangular, 75ST6 aluminumalloy plate by nonuniform heating are determined both experimentally and theoretically. The characteristics of commercially available bond ed resistance wire strain gages are first investigated to determine their suitability for measuring stresses under simple conditions of stress and temperature. The gages are then used to measure thermal stresses in the flat plate in order to study their suitability under more complicated conditions. The experi mental results are found to be in satisfactory agree ment (within t5 percent of maximum calculated stress) with an approximate theoretical solution of the problem. NACA RESEARCH ABSTRACTS NO.29 NACA TN 2770 STUDY OF THE PRESSURE RISE ACROSS SHOCK WAVES REQUIRED TO SEPARATE LAMINAR AND TLRRBLiLENT BOUNDARY LAYERS. ColemanduP. Donaldson and Roy H. Lange. Stptenetr 1952. 20p. diagrs., photos., tab. (NACATN 2770. Formerly RM L52C21) Results are presented of a dimensional study and an experimental investigation of the pressure rise across a shock wave which causes separation of the boundary layer on a flat plate. The experimental part of the investigation was conducted at a Mach number of 3. 03 for a Reynolds number range of 2 x 106 to 19 x 106. The available experimental data are compared with the predictions of the present studi,., and the >iLiifi in, e of the results obtained is discussed relative to certain practical design prob lems. NACA TN 2771 THERMAL BUCKLING OF PLATES. Myron L. Gossard, Paul Seide, and William M. Roberts. August 1952. 39p. diagrs. (NACA TN 2771) An approximate method, based on largedeflection plate theory, for calculating the deflections of flat or initially imperfect plates subject to thermal buckling is outlined. The method is used to determine the de flections of a simply supported panel subjected to a tentlike temperature distribution over the plate sur face. Experimental results for a particular panel are in good agreement with the theoretical results considered in the test. NACA TN 2772 DRIVING STANDING WAVES BY HEAT ADDITION. Perry L. Blackshear, Jr. August 1952. 46p. diagrs., photos. (NACA TN 2772) Types of burner instability are enumerated and the role of standing waves in burners is discussed. The status of the problem of flamedriven standing waves is reviewed and a onedimensional flow theory giving the mechanism whereby a flame drives or damps a standing wave is presented. In this theory, the re flection, transmission, and amplification of waves passing through a flame region were determined from the continuity and momentum equations. For the model considered, waves were found to pass through the flame front with their velocity amplitudes unal tered so long as the flame area remained unchanged. A change in flame area acted as a source of waves propagating simultaneously into the hot and cold gases on either side of the flame zone. The one dimension al theory seems an adequate explanation of the exper imental observations. NACA RESEARCH ABSTRACTS NO.29 NACA TN 2774 A METHOD FOR FINDING A LEASTSQUARES POLYNOMIAL THAT PASSES THROUGH A SPECI FIED POINT WITH SPECIFIED DERIVATIVES. Neal Tetervin. September 1952. lip. (NACA TN 2774) A method is presented for finding a mthdegree poly nomial that passes through a specified point with (m j) specified derivatives (1 = j < m) and is a leastsquares polynomial for other points which are spaced at unequal intervals of the independent vari able. NACA TN 2775 EFFECT OF LINEAR SPANWISE VARIATIONS OF TWIST AND CIRCULARARC CAMBER ON LOW SPEED STATIC STABILITY, ROLLING, AND YAW ING CHARACTERISTICS OF A 45 SWEPTBACK WING OF ASPECT RATIO 4 AND TAPER RATIO 0.6. Byron M. Jaquet. August 1952. 27p. diagrs., 2 tabs. (NACA TN 2775) An investigation was made in the Langley stability tunnel at low scale to determine lowspeed effects of linear spanwise variations of twist and circulararc camber on staticstability and rotarystability (roll ing and yawing) derivatives of a 45 sweptback wing of aspect ratio 4 and taper ratio 0.6. NACA TN 2776 THE EFFECT OF A SIMULATED PROPELLER SLIPSTREAM ON THE AERODYNAMIC CHARAC TERISTICS OF AN UNSWEPT WING PANEL WITH AND WITHOUT NACELLES AT MACH NUMBERS FROM 0. 30 TO 0. 86. Gareth H. Jordan and Richard I. Cole. September 1952. 15p. diagrs., photo. (NACA TN 2776) Results of an investigation are presented to show the effect of a simulated propeller slipstream on the lift, drag, and pitchingmoment characteristics of an un swept wing panel with and without nacelles. The data were obtained in the Langley 24inch highspeed tunnel through a range of Mach numbers from ap proximately 0. 30 to 0. 86 for angles of attack of 0 and 3. NACA TN 2777 THEORETICAL DISTRIBUTION OF SLIP ANGLES IN AN AGGREGATE OF FACECENTERED CUBIC CRYSTALS. John M. Hedgepeth. August 1952. 32p. diagrs. (NACA TN 2777) An analysis of the relative frequency of occurrence of any given slipline angle in a plastically deformed polycrystal composed of facecentered cubic crystals is presented for the case of simple tension. The re sults are compared with those obtained for a poly crystal composed of crystals which have but a single 3 mode of slip and with experimental results. The comparisons show that the differences between the results obtained by the two theories become greater as the stress is increased. The comparison with experiment of the facecentered cubic theory is somewhat better than that of the singleslipmode theory, but the errors are appreciable. NACA TN 2778 STRESS AND STRAIN AT ONSET OF CRAZNLu OF POLYMETHYL METHACRYLATE AT VARIOUS TEMPERATURES. M. A. Sherman and B. M. Axilrod, National Bureau of Standards. September 1952. 21p. diagrs., photos., 3 tabs. (NACA TN 2778) Stress and strain at the onset of crazing were deter mined for both generalpurpose and heatresistant grades of commercial cast polymethyl methacrylate. Tests were made at 230, 50, and 70 C on samples 0. 15 inch thick. Other properties measured were tensile ti enth. total elongation, and modulus of elasticity. Results obtained indicate that a "critical strain theory" for the threshold of crazing is not ap plicable to polymethyl methacrylate. Strain at the onset of crazing decreased with increase in tempera ture from 23 to 50 C whereas there was no consist ent trend from 50 to 70 C. Stress at the onset of crazing was about 80 to 95 percent of the tensile strength at all three temperatures. NACA RM E52F19 VISUALIZATION OF SECONDARYFLOW PHENOME NA IN BLADE ROW. H. Z. Herzig, A. G. Hansen and G. R. Costello. August 1952. 20p. photos., diagrs. (NACA RM E52F19) Flowvisualization methods were applied in a prelim inary study of the streamline pattern of secondary flows in a blade row. The investigation demonstrated flow of the inletwall boundary layer in a blade pas sage into the corner between the blade suction surface and the wall to form a vortex. This vortex formation occurs well within the passage and is not a trailing edge phenomenon. The magnitudes of the velocity gradients and angle deflections make the significance of quantitative measurements in this region question able. NACA RM E52G17 EFFECT OF FREE METHYL RADICALS ON SLOW OXIDATION OF PROPANE AND ETHANE. Glen E. McDonald and Rose L. Schalla. August 1952. 21p. diagrs., 4 tabs. (NACA RM E52G17) A study of the effect of free methyl radicals on the slow oxidation of both ethane and propane was made by means of a photochemical decomposition reaction in a static system. The introduction of methyl radi cals into a mixture of propane and oxygen or ethane and oxygen substantially lowered the initiation tem 4 perature of the combustion of the hydrocarbon. Measurements were made of the amount of propane oxidized in the presence of free methyl radicals at temperatures ranging from 0 to 116 C and for two different concentrations of propane. Oxidation of ethane, which was induced by the free nietii.l radical, was investigated for only one concentration from 0 to 183 C. From a plot of the log percentage of propane oxidized against the reciprocal temperature (OK), a zero activation energy, was found for the reaction be low 400 C. For ethane the activation eiieriV was zero below 120 C. Above these two temperatures an increase in the activation energy was observed. A mechanism is proposed to account for the zero acti vation enertr",, but no attempt was made to interpret the .iihc temperature reactions. NACA RM E52G24 INFLUENCE OF EXTERNAL VARIABLES ON SMOK ING OF BENZENE FLAMES. Thomas P. Clark. August 1952. 17p. diagrs. (NACA RM E52G24) Premixed benzeneair flames were burned with vari ations in initial gas temperature, fuelflow rate, flame length, secondaryairflow rate and burner tube diameter. Fuelair ratios were measured both at the initial appearance of incandescent carbon in the flame and at the point where smoke issued from the flame. Temperature alone affected the fuelair ratio at which carbon first formed in the flame. The flame did not smoke under a fuelair ratio of 164 per cent of stoichiometric, id ,lltss of the environment around the flame. Conditions which enhanced the dif fusion of ,:,.,'t into the outer cone improved the smokeburning characteristics of the flame. NACA TM 1343 THE EXCITATION OF UNSTABLE PERTURBATIONS IN A LAMINAR FRICTION LAYER. (Die Anfachung instabiler Storungen in einer laminaren Reibungs schicht). J. Pretsch. September 1952. 63p. diagrs., 3 tabs. (NACA TM 1343. Trans. from Aerodynamische Versuchsanstalt Gottingen E. V., Institute fur Forschungsflugbetrieb und Flugwesen; Jahrbuch der deutschen Luftfahrtforschung, August 1942, p. I 5471). With the aid of the method of small oscillations which was used successfully in the investigation of the stability of laminar velocity distributions in the presence of twodimensional perturbations, the excitation of the unstable perturbations for the Hartree velocity distributions occurring in plane boundary layer flow for decreasing and increasing pressure is calculated as a supplement to a former report. The results of this investigation are to make a contribu tion toward calculation of the transition point on cylindrical bodies. NACA TM 1348 ON THE REPRESENTATION OF THE STABILITY REGION IN OSCILLATION PROBLEMS WITH THE AID OF THE HURWITZ DETERMINANTS. (Zur NACA RESEARCH ABSTRACTS N0.29 Darstellung des Stabilitatsgebietes bei Schwingungs aufgaben mit Hilfe der HurwitzDeterminanten). E. Sponder. August 1952. 12p. diagrs. (NACA TM 1348. Trans from Schweizer Archiv fur ange wandte Wissenschaft und Technik, v. 16, no. 3, March 1950, p. 9396). This report concerns the use of the Hurwitz deter minants in defining boundaries of regions where os cillatory phenomena are to be stable or unstable. A simplification is suggested as an aid in reducing the computations usually required, although it is empha sized that point checks in the various regions defined are required using the complete set of Hurwitz de terminants or some other complete stability deter mination. NACA TM 1352 TRANSITION CAUSED BY THE LAMINAR FLOW SEPARATION. (S6ryuHakuri ni tomnionau Sen'i ni kansuru Kenkyu). T. Maekawa and S. Atsumi. September 1952. 14p. diagrs., 2 tabs. (NACA TM 1352. Trans from Society of Applied Mechanics of Japan, Journal, v. 1, no. 6, November 1948, p. 187192) An experimental investigation of the effects of the geometry of body surface, Reynolds number, stream turbulence, and a roughness element (wire) on the reattachment of separated laminar boundarylayer flow on a bent flat plate is presented and discussed. The flow mechanisms determining reattachment of the boundary layer are analyzed and discussed. BRITISH REPORTS N16809* Aeronautical Research Council (Gt. Brit.) GROUND EFFECT ON DOWNWASH WITH SLIP STREAM. P. R. Owen and H. Hogg. 1952. 12p. diagrs., 8 tabs. (ARC R M 2449; ARC 7590. Formerly RAE Aero 1901) It is shown in this report that given the downwash at the tail plane with slipstream, then the decrement in downwash at the same incidence and propeller thrust coefficient due to ground interference can be esti mated by the formula of Piper and Davies (with cer tain slight modifications) with sufficient accuracy for most practical purposes. N16810* Aeronautical Research Council (Gt. Brit.) TESTS ON A CONTROL WITH FIXED TRAILING EDGE ANGLE, AND SURFACES OF VARIABLE CAMBER. A. S. Halliday, H. Deacon and W. C. Skelton. 1952. 15p. diagrs., 3 tabs. (ARC R & M 2495. Formerly ARC 10, 598; S &C 2116) NACA RESEARCH ABSTRACTS NO.29 The report describes experiments to determine the effect of surface distortion on the hinge moments of a control and to investigate the importance of a rigid trailing edge in the prevention of large changes in hingemoment coefficient. The results show that by designing the control so that the surface at the trail ing edge, extending over oneeighth of the chord, is incapable of distortion, then the value of b2 remains almost constant within the limits on any possible dis tortion forward of the fixed portion. It has also been found that bI is more constant. Changes in center line camber of the deformed portion of the control can still produce an effective change in b2 when the cam ber varies with the control angle, although the amount of this change may be roughly halved. Present prac tice of treating trailingedge angle as an important parameter in specifying control characteristics is well justified. N16811'* Aeronautical Research Council (Gt. Brit.) LIMITATIONS OF USE OF BUSEMANN'S SECOND ORDER SUPERSONIC AEROFOIL THEORY. W. F. Hilton. 1952. 9p. diagrs. (ARC R & M 2524. Formerly ARC 9869; FM 971) The author has found the Busemann theory very rapid in use for the determination of CL, CD, and Cm, and this report will enable the exact scope of its use to be determined. It has been tacitly assumed in the past that Busemann's secondorder theory of airfoils at supersonic speeds was subject to the same limita tions of wedge angle as the exact theory given by Lighthill and others, namely, the wedge angle at which the bow wave detaches. The range of angles for which Busemann's theory gives a pressure coef ficient in error by less than 1 percent is shown to be smaller than the angle range for the shock wave to be attached. There is also a limit to the application of Busemann's method to angles of expansion as well as to angles of compression, unlike the exact theory, which can be extended to expansive angles of the order of one right angle without breaking down, in fact far beyond the useful range. The limits of angle given for the use of Busemann's theory are conserv ative, since they give the pressures to 1 percent, and the force coefficients will be more accurately de termined since the errors tend to cancel out when integrating pressures to obtain forces. N16812* Aeronautical Research Council (Gt. Brit.) ABSTRACTS OF PAPERS PUBLISHED EXTERN ALLY. 1952. 17p. (ARC R & M 2565) Presents an index and related abstracts of papers discussed by the Aeronautical Research Council and recommended by them for external publication. 5 N16813" Aeronautical Research Council (Gt. Brit.) NOTE ON THE EFFECT OF SIZE OF AIRCRAFT UPON THE DIFFICULTIES INVOLVED IN LANDING AN AIRCRAFT. D. Adamson. 1951. 13p. diagrs., tab. (ARC R & M 2567; ARC 10,903. Formerly RAE Aero 2202) The effect of aircraft size upon the response of an aircraft during those maneuvers which are commonly employed in landing has been examined, and in this way an assessment has been made of the way in vhich the difficulties experienced by the pilot will go up as aircraft size increases. On the basis of the work summarized in this note, it is concluded that the problems associated with landing (from the pilot's point of view at any rate) are unlikely to be aggrava ted to such an extent, as the size of aircraft increases up to the limiting size considered in this report (300 ft span), as to make landing really difficult. N16814* Aeronautical Research Council (Gt. Brit.) ON THE MOMENTUM EQUATION IN LAMINAR BOUNDARYLAYER FLOW. A NEW METHOD OF UNIPARAMETRIC CALCULATION. B. Thwaites. 1952. 9p. (ARC R & M 2587. Formerly ARC 11,155; FM 1193; Perf. 397) The general method of Pohlhausen, discussed in de tail in reference 1, uses a uniparametric system of velocity distributions of the form u/U = f(y/ 6) + xg(y/5). Pohlhausen, by choosing simple forms for the functions f and g, then uses the momentum equation to find the distribution of 6 with x and thence the distributions with x of the other bound ary layer characteristics. Several awkwardnesses exist in his method, especially when it is applied to problems dealing with a normal velocity at the bound ary. In this paper, a new method is described of combining velocity distributions in the form y/O = F(u/U) + AG(u/U), and it is shown that such a combination avoids several difficulties. This method of combination also allows a second parameter apart from X, which might be found valuable in certain problems. The method has been briefly described before as part of an investigation into the effect of continuous suction on laminar boundarylayer flow under adverse pressure gradients. In that paper (R & M 2514), a numerical example of its use was given. In this paper, no example will be given be cause, as far as the author can see, the practical use of the method is superseded by the generalized method of reference 1; however, it possessed con siderable analytical interest. N16815* Aeronautical Research Council (Gt. Brit.) COMPARATIVE TESTS OF THICK AND THIN TURN ING VANES IN THE ROYAL AIRCRAFT ESTABLISH MENT 4 x 3FT WIND TUNNEL. K. G. Winter. 1952. 5p. diagrs., tab. (ARC R & M 2589; ARC 10,976. Formerly RAE Aero 2217) 6 The tests were made by replacing the existing center six thick vanes at the first corner of the 4 x 3ft wind tunnel by vanes of sheet metal. The thin vanes reduced the corner loss, estimated from a wake traverse behind one vane, without any deterioration in outflow, and are therefore recommended for use in future wind tunnels. N16816* Aeronautical Research Council (Gt. Brit.) TABLES OF MULTHOPP AND OTHER FUNCTIONS FOR USE IN LIFTINGLINE AND LIFTINGPLANE THEORY. V. M. Falkner. WITH APPENDIX: CALCULATION OF THE CIRCULATION FROM THE DOWNWASH. E. J. Watson. 1952. 52p. diagrs., 8 tabs. (ARC R & M 2593. Formerly ARC 11, 234; S & C 2187; Perf. 405) The report gives the derivation and computed tables of two classes of functions suitable for the solution of problems of spanwise aerodynamic loading of wings either by liftingline or liftingplane theory. The functions are based on liftingline theory, but, by a consideration of the connection between lifting line and liftingplane theory through the application of Munk's stitit: r theorem to the calculation of in dueed (I rai. it is deduced that the functions must be equally suitable for liftingplane theory. The first range of functions, called Multhopp or M functions, is associated with discontinuities of induced downwash, while the second, called P functions because of the p.,1%. v ial representation of induced downwash, is connected with discontinuities in rate of change of induced downwash. Examples are given of the com bination of functions to produce given curves of in duced downwash, and evidence of the close relation between the results for a continuous and stepped downwash curve suggests that the functions tabulated will be sufficient to cover almost any problem in wing loading. N16817* Aeronautical Research Council (Gt. Brit.) CALCULATED AERODYNAMIC CHARACTERISTICS OF TWO INFINITE WINGS WITH CONSTANT CHORD. V. M. Falkner. 1952. lOp. diagrs., 8 tabs. (ARC R & M 2594. Formerly ARC 10,628; S & C 2123; Perf. 325) The report gives solutions obtained by the vortex lat tice method for the aerodynamic loading of two infi nite wings of constant chord with sweepback of 45, one with a Vjoint at the center, the other rounded off with arcs of radius four times the chord. The true mathematical solution for these problems is exceed ingly difficult to find, and the accuracy has been veri fied by considering the convergence of solutions of varying complexity. The Vwing shows a reduction in circulation near the joint with accompanying back ward movement of the local center of pressure, while the rounded wing has increased circulation without appreciable variation of the center of pressure from the 0. 25chord position. The results will be used to NACA RESEARCH ABSTRACTS NO.29 modify loading functions used in vortex lattice theory in order to improve solutions for wings of small as pect ratio, particularly when the leading or trailing edges meet at an included angle which differs consid erably from 180. N1681b* Aeronautical Research Council (Gt. Brit.) APPLICATION OF THE LINEAR PERTURBATION THEORY TO COMPRESSIBLE FLOW ABOUT BODIES OF REVOLUTION. A. D. Young and S. Kirkby. 1952. 8p. diagrs. (ARC R & M 2624; ARC 11,033; FM 1177. Formerly College of Aeronautics, Cran field Rept. 11) The linearized theory is developed in some detail in order to clarify the difference between two dimension al flow and flow past slender threedimensional bod ies. In agreement with other authors, it is conclud ed that the perturbation velocity on a slender body in compressible flow is 1/'62 times the perturbation ve locity in incompressible flow on a thinner body at re duced incidence obtained by reducing the lateral di mensions of the original body in the ratio 3:1. This result is applied to a representative family of stream line bodies of revolution at zero incidence. It is found that, without an undue loss of accuracy, the re sults of the claculations can be presented in a rela tively simple form in a diagram showing the variation of velocity with Mach number for a range of values of velocity on the surface of a streamline body in incom pressible flow. This variation is always less than that predicted by the Glauert law but approaches it with increase in the basic incompressible flow veloc ity, being very close to it for basic incompressible velocity ratios, u/U0, of 1.10 and higher. It is shown that the blockage factor for a body of revolu tion in a wind tunnel is increased in compressible flow in the ratio 1/33 and not 1/i34 as quoted in R &M 1909. N16819* Aeronautical Research Council (Gt. Brit.) MODEL TESTS IN THE 24FT WIND TUNNEL TO DETERMINE THE OPTIMUM ANGLE FOR RELEASE OF A COCKPIT HOOD. R. Fail. 1952. 3p. diagrs. (ARC R & M 2644; ARC 11, 494. Formerly RAE Tech. Note Aero 1947) For some time now, it has been recommended that mechanical assistance be incorporated in jettisonable cockpit hood designs. Some firms have preferred designs in which the hood is constrained to rotate through a definite angle before the final release. A short series of tests has, therefore, been made in the 24ft wind tunnel to determine the optimum angle for release. This was found to be about 10, which is considerably less than has been suggested in the past. N16820* Aeronautical Research Council (Gt. Brit.) WINDTUNNEL TESTS ON A THICK SUCTION AERO FOIL WITH A SINGLE SLOT. M. B. Glauert, W. S. Walker, W. G. Raymer and N. Gregory. 1952. NACA RESEARCH ABSTRACTS NO.29 14p. diagrs., tab. (ARC R &M 2646; ARC 11,797. Formerly ARC 10, 854; Perf. 348; FM 1150) This report describes the twodimensional wind tunnel tests carried out in the National Physical Lab oratory 13 by 9ft wind tunnel on a 31. 5percent thick suction airfoil, GLAS II, which has a single slot on the upper surface at 69 percent chord. Both suc tion and blowing were used to prevent separation. Lift, drag, pitching moment, and the flow through the slot were measured. Tests without suction were made at Reynolds numbers of 0. 96 and 2. 8b million. The results at the two Reynolds numbers were mark edly different, and at the higher speed widely varying values of the drag coefficient were recorded in the same conditions, there apparently being several pos sible regimes of flow. With suction, the pump power available only enabled tests to be made at the lower Reynolds number, and with the boundary layer on the upper surface laminar to the slot. At low incidences, suction quantities agreeing well with theoretical esti mates sufficed to maintain unseparated flow, but at higher incidences the flow tended to break down. Three or four times as much suction was required at all incidences to make the separated flow readhere. With blowing, still larger quantities were necessary, but the spanwise distribution of the flow from the slot was unsatisfactory. Two different slot shapes were tested on the model, one with a sharp beak to the front lip, the other with a rounded entry. Intermit tent separation of the flow occurred in each case. The phenomena may be of a fundamental character and associated with the profile shape rather than with the shape of the slot entry. N16947C* Aeronautical Research Council (Gt. Brit.) CONTRIBUTIONS TO UNSTEADY AEROFOIL THEORY. IX. A NUMERICAL COMPARISON OF VARIOUS APPROXIMATE THEORIES FOR THE OSCILLATING AEROFOIL OF LARGE SPAN. Katterbach. August 24, 1948. 27p. diagrs., 6 tabs. (ARC 11,736; ARC 0. 755) The lift coefficients of an oscillating elliptical wing for the reduced frequencies Q = ivb/v = 0, 0. 5, 1.0, ... 6. 5, are calculated according to four well known approximate theories which deal with the un steady motion of an airfoil; the results, expressed in complex form, are compared with each other and with strip theory. The theories are described brief ly, and the equations required for the calculation are derived. N16967* Aeronautical Inspection Directorate (Gt. Brit.) MECHANICAL THUMB HARD DRYING TIME TEST ER FOR PAINTS. A. T. Josling. May 1952. 9p. photos., diagrs. (AID Rept. APP. 2) An original mechanical thumb drying time tester, capable of reproducing results with varying types of paints, mention being made of protection coating and other metal finishes. 7 N16968* Aeronautical Inspection Directorate (Gt. Brit.) APPARATUS FOR OBTAINING A WIDE RANGE OF HUMIDITY AND TEMPERATURE CONDITIONS IN A STREAM OF AIR. A. T. Josling. May 1952. 4p. diagrs., photos. (AID Rept. APP. 3) A new type of apparatus for obtaining any desired humidity in a stream of air is described. It is based on the principle of saturating air at a suitable tem perature and then heating it to the working tempera ture. Saturation is obtained by means of a special splashing device without atomization. Since the sat uration temperature to give the desired humidity at the working temperature, which may be i '.%,1 1,. within its wide range, is known from ti',,,. metricc tables, an actual determination of the humidity achieved is rendered unnecessary. The apparatus is suitable for blushing tests on paints and dopes, cor rosion tests on platings, protective coatings, oils and greases. N169694 Aeronautical Inspection Directorate (Gt. Brit.) DEVICE FOR COLD BEND TESTS ON PAINT TEST PANELS. A. T. Josling. May 1952. 3p. diagrs. (AID Rept. APP. 4) A device enabling bend tests to be performed at cor rect temperatures. A hinged test panel holder is also shown. N16970* Aeronautical Inspection Directorate (Gt. Brit.) THE MEASUREMENT OF THE WALL THICKNESS OF A LONG TUBE. H. R. Davenport and E. W. Carter. December 5, 1951. 7p. diagrs., photos. (AID Rept. Metro 4) The description of a method of measuring the wall thickness of a long tube by means of an alinement telescope and auxiliary equipment. N17019* Forest Products Research Lab. (Gt. Brit.) TRIALS OF TIMBER FOR PLYWOOD MANUFAC TURE. PROGRESS REPORT SIXTEEN. June 1952. 9p. 5 tabs. (Forest Products Research Lab.) Tests were made on three British Guiana species of wood. They were crabwood (carapa guianensis), kurokai (protium decandrum), and wamara (swarzia leiocalycina). It was found that the crabwood would be unsuitable for British mills but it may be of some use for conversion in the country of origin if freshly felled logs are less prone to splitting and shaking. Kurokai would also be unsuited to British mills but would be of use in the country or origin. Wamara is considered unsuitable because of its weight and pro cessing difficulties. Uilivcmot I Vr Jr Li.if\Lht^f NACA 3 1262 09079 7472 8 RESEARCH ABSTRACTS NO. MISCELLANEOUS NACA Rept. 1037 Errata No. 1 on "GENERAL METHOD AND THER MODYNAMIC TABLES FOR COMPUTATION OF EQUILIBRIUM COMPOSITION AND TEMPERATURE OF CHEMICAL REACTIONS". Vearl N. Huff, Sanford Gordon and Virginia E. Morrell. 1951. NACA TN 2277 Errata No. 1 on "EFFECTS OF COMPRESSIBILITY ON THE PERFORMANCE OF TWO FULLSCALE HELICOPTER ROTORS'. Paul J. Carpenter. January 1951. NACA TN 2353 Errata No. 1 on "CHARTS AND TABLES FOR USE IN CALCULATIONS OF DOWNWASH OF WINGS OF ARBITRARY PLAN FORM'. Franklin W. Diederich. May 1951. UNPUBLISHED PAPERS N5292* FORMS OF WING FLUTTER IN TWODIMENSIONAL COMPRESSIBLE POTENTIAL FLOW. (Flugelsch wingungsformen in ebener kompressibler Potential strcmung). Nikolaus Rott. July 1952. 44p. diagrs. (Trans. from Zeitschrift fur angewandte Mathematik und Physik, v. 1, 1950, p. 380410) On the basis of energy considerations, a survey is given of the possible forms of oscillations for flutter with two degrees of freedom in a plane compressible potential flow. It is found that the possible forms do not much depend on Mach number in the whole range from 0 to a, with some exceptions in the neighbor hood of M = 1, where forms with only one degree of freedom (torsion) may occur. Conclusions are drawn for methods of preventing flutter. Limits of the reduced frequency are given for the possibility of flutter in two and one degree of freedom. Special care was given for the case of M = 1. NACALangley 91652 4000 
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