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fNational Advisory Committee for Aeronautics i:a: Research Abstracts iO.S FEBRUARY20, 1953 :CURRENT NACA SACIA Rept. 1069 N SOLUTIONN OF THE NONLINN 'L EQUATION FOR TRANSONIC ;O PAST A tAVESHAPED WALL. Carl Kapla j iA. , :.lp. diagrs., tab. (NACA Rept. 1069. For  TS:383) h e simplified nonlinear differential equation for trao~lc flow past a wavy wall is solved by the method of integration iti series. The solution has ui:been carried to the point where'the question of the tftence dr rdonexistence of a mixed potential flow icai be answered by the behavior of a single power ~ selies in the transonic similarity parameter. The a, culation of the coefficients of this dominant power alries has been reduced to a routine computing protdinm by means of recursion formulas resulting rIlom the solution of the differential equation and the i~.1andary condition at the surface of the wavy wall. VNAA Rept. 1071 : THEORETICAL SYMMETRIC SPAN LOADING DUE TO FLAP DEFLECTION FOR WINGS OF ARBITRARY PLAN FORM AT SUBSONIC SPEEDS. John DeYoung. 952. ii, 41p. diagrs., tabs. (NACA Rept. 1071. Formerly NACA TN 2278) p .' l cedure based upon a simplified liftingsurface :; theory that includes effects of compressibility and % !j:.::panwise variation of section liftcurve slope is pre Scented in such a manner that the spanwise loading due Stoflap deflection can be simply found for wings hav P. :j symmetric plan forms with constant spanwise i'.~iaweep angle of the quarterchord line. Aerodynamic characteristics due to flap deflection are considered k. .la sidlor straighttapered wings, values of certain of ': these characteristics are presented in charts for a Z;; range of swept plan forms. Further use of the meth od gives downwash in the vertical center of the wake !: of the wing. 4' ; .ACA Rept. 1083 ': AXJSYMMETRIC SUPERSONIC FLOW IN ROTATING ; IMPELLERS. Arthur W. Goldstein. 1952. ii, ., 14p, diagrs. (NACA Rept. 1083. Formerly T. 2388) General equations are developed for isentropic, f,: .rictionless, scisymmetric cbmpressible flow in S rotating impellers with blade forces eliminated in : r"favor of the bladesutface function. The character i., tic equations for supersonic flow are developed 88LE ON LOAN ONLY. IR UESTS FOR DOCUMENTS TO NACA, 1794 F ST., NW., )RT Ti LE AND AUTHOR. t'i ". a c;'.. ." .!*.. and a computing technique is utilized to find the ef fect of variations of design parameters on internal flow and workinput distribution. NACA TN 2881 AERODYNAMIC CHARACTERISTICS OF A TWO BLADE NACA 10(3)(062)045 PROPELLER AND OF A TWOBLADE NACA 10(3)(08)045 PROPELLER. William Solomon. January 1953. 53p. diagrs., photo., tab. (NACA TN 2881. Formerly RM L8E26) Characteristics are given for the twoblade NACA 10(3)(062)045 propeller and for the twoblade NACA 10(3)(08)045 propeller over a range of ad vance ratio from 0.5 to 3.8, through a bladeangle range from 200 to 550 measured at the 0.75 radius. Maximum efficiencies of the order of 91.5 to 92 per cent were obtained for the propellers. The propeller with the thinner airfoil sections over the outboard portion of the blades, the NACA 10(3)(062)045 pro peller, had lower losses at high tip speeds, the difference amounting to about 5 percent at a helical tip Mach number of 1.10. NACA TN 2884 CALCULATION AND MEASUREMENT OF NORMAL MODES OF VIBRATION OF AN ALUMINUMALLOY BOX BEAM WITH AND WITHOUT LARGE DIS CONTINUITIES. Frank C. Smith and Darnley M. Howard, National Bureau of Standards. January 1953. 40p. diagrs., photo., 8 tabs. (NACA TN 2884) The lowest normal modes of vibration of three alumi num alloy box beams were calculated using a matrix iteration method. For the calculations the actual structures were idealized to a system of mass points interconnected by massless springs. The lowest normal modes ol these beams were measured experi mentally and compared with those calculated. This comparison indicates that the mode shapes and natural frequencies for structures of this type may be adequately calculated using this method. The experi mental measurements were limited at the higher frequencies by local vibrations of small elements of the beams. NACA TN 2885 SOME EXACT SOLUTIONS OF TWODIMENSIONAL FLOWS OF COMPRESSIBLE FLUID WITH HODO GRAPH METHOD. ChiehChien Chang and Vivian O'Brien, Johns Hopkins University. February 1953. 63p. diagrs., 4 tabs. (NACA TN 2885) WASHINGTON 95. D. C, CITING CODE NUMBER ABOVE EACH TITLE; 2 A suggestion is given for classifying compressible potential flows according to location and number of singularities in the subsonic region of the hodograph plane, which seems to offer a convenient, criterion for systematic investigation of these flows with Chaplygin's original method. The object of the pa per is to present and analyze a few useful solutions of compressible potential flow with the exact gas law. These solutions include flows about convex corners and belong to the same class as that of Ringleb. Also, the exact solution of compressible flow through a particular contracting channel is given. NACA TN 2886 AN ANALYSIS OF STATICALLY INDETERMINATE TRUSSES HAVING MEMBERS STRESSED BEYOND THE PROPORTIONAL LIMIT. Thomas W. Wilder, II. February 1953. 13p. diagrs., 4 tabs. (NACA TN 2886) A procedure for analyzing statically indeterminate trusses in the plastic stress range is presented which is applicable to trusses having any number of redun dant members. By using the RambergOsgood ana lytical representation of the stressstrain curve, the analysis of the truss is reduced to the solution of a set of simultaneous equations. A numerical example is presented to illustrate the procedure. NACA TN 2888 PERFORMANCE CHARACTERISTICS OF PLANE WALL TWODIMENSIONAL DIFFUSERS. Elliott G. Reid, Stanford University. February 1953. 1, 80p. diagrs., photos., 3 tabs. (NACA TN 2888) Performance characteristics were determined for planewall, twodimensional diffusers which were so proportioned as to insure reasonable approximation of twodimensional flow. The diffusers had identical entrance cross sections and discharged directly into a large plenum chamber; the test program included wide variations of divergence angle and length. A dynamic pressure of 60 pounds per square loot was maintained at the diffuser entrance and the boundary layer there was thin and fully turbulent. A few tests were made with asymmetric diffusers. Others showed the effects of addition of a short exit duct of uniform section and of installation of a thin, central, longitudinal partition. NACA TN 2891 FACTORS AFFECTING LAMINAR BOUNDARY LAY ER MEASUREMENTS IN A SUPERSONIC STREAM Robert E. Blue and George M. Low. Appendix B: REDUCTION OF DATA. Jack M Lande. February 1953. 49p. diagrs. (NACA TN 2891) The observed discrepancy at supersonic speeds be tween theoretical and apparent experimental average flat plate Irictiondrag coefficients calculated from boundary layer totalpressure surveys was investi gated. Effects of the totalpressure probe, heat transfer through the leadingedge region, leading NACA RESEARCH ABSTRACTS NOiI :' .. edge geometry and strength of the leadlmgedfle, possible early transition to turbulent flow drb "b of turbulence, and the slight streamwil tift gradient inherent in flatplate flow were i li the investigation. Only one of these fact, 'tfhe:. feet of the totalpressure probe, was foudti..e ... significant. Totalpressure probes of different tip heights, when placed in laminar boundary la yeiAe veloping under identical conditions, measure dz er ent values of frictiondrag coefficient. EG"rapolk, tion of these measurements indicates that sa .*u of vanishing tip height would measure the theoreticall;y predicated values of average flat plateffrictiond ig coefficients. NACA TN 2894 CALCULATIONS OF UPWASH IN THE REGION ABOVE OR BELOW THE WINGCHORD PLANESOF. . SWEPTBACK WINGFUSELAGENACEL.LE. COM BINATIONS. Vernon L Rogallo and.JohnL. ' McCloud, m. February 1953. 15p. diagra, photo".' (NACA TN 2894) A procedure has been developed for predicting theW.. upwash components of the upflow angles in the reglt.i. above or below the wingchord planes of sweptback .ifl wingfuselagenacelle combinations. Compariso.ns are made of the predicted and measured upflow an :. gles for six semispan models with 4Q sweptback ' wings. \ NACA TN 2896 SURVEY OF PORTIONS OF THE IRONNICKEL . MOLYBDENUM AND COBALTIRONMOLYBDEN IM TERNARY SYSTEMS AT 12000 C. Dilip K. Da an d': Paul A. Beck, University of Notre Dame. :' February 1953. 56p. diagrs., photos., I& tabs. :: (NACA TN 2896) The 12000 C isothermal sections of the ironnickel molybdenum and the cobaltironmolybtenuni teparay systems were surveyed. The phases occurringn. these systems were identified by means ot Xray' diffraction and by etching methods, and the phase' boundaries at 12000 C were determined microlIeopi'' cally, using the disappearing phase method with' ' quenched specimens. Both systems contain long : solidsolution fields of the mu phase. Other inter . mediate phases occurring in the ironnickel :: molybdenum system are the P phase and the delta phase. Both phase diagrams have extensive face centered cubic solidsolution fields and some body centered cubic solid solutions. NACA RM E52L09 FORCEDCONVECTION HEATTRANSFER CHAR . ACTERISTICS OF MOLTEN SODIUM HYDROXIDE ... Milton D. Grele and Louis Gedeon. February. 1953,., 27p. diagrs., photo., 2 tabs. (NACA RM E52LO.B). .. The forcedconvection heattransfer characteristics :: of sodium hydroxide were experimentally investi gated. The heattransfer data for heating fall slghtl '. .NACA RESEARCH ABSTRACTS NO.38 ly above the McAdams correlation line, and the heat transfer data for cooling are fairly well represented by the McAdams correlation line. I qACA TM 1344 ION THE THEORY OF THE TURBULENT BOUNDARY .AYER. (Uber die Theorie der turbulenten i' renzachichten). J. Rotta. February 1953. 50p. fagrs. (NACA TM 1344. Trans. from Max 'latickInstitut fTr Stromungslorschung, G6ttingen. i1tteilungen 1, 1950) turbulent energy, dissipation, and momentum rela Stionsare discussed. A procedure is given for com puttation of turbulent skin friction in boundarylayer .flow ith pressure gradients. The boundary layer is divided into a region near the wall where viscosity and surface roughness are important, an outer re gion which is dependent on friction coefficient and pressure gradient, and an intermediate zone between these two which is unaffected by wail roughness, vis "cosily, and the outer flow. Analytical confirmation is obtained for the empirical fact that turbulent boundary layers are able to overcome a greater pressure rise than laminar ones. NACA TM 1349 ON A CLASS OF EXACT SOLUTIONS OF THE EQUATIONS OF MOTION OF A VISCOUS FLUID. (Ob odnom klasse tochnykh reshenii uravnenii dvizheniya vyazkoi zhidkosti). V. I. Yatseyev February 1953. 7p. (NACA TM 1349. Trans. from Zhurnal Eksperimental 'noi i Teoretisheskoi Fiziki, v.20, no. 11, 1950, p. 10311034). Thie general solution is obtained of the equations of motion of a viscous fluid in which the velocity field is.inversely proportional to the distance from a certain point. Some particular cases of such mo tion are investigated. NACA TM 1358 CALCULATION OF THE SHAPE OF A TWO DIMENSIONAL SUPERSONIC NOZZLE IN CLOSED FORM. (Sul Calcolo in Termiru Finiti dell'Effusore di una Galleria Bidimensionale Supersonica). Dante Cunsolo. January 1953. 29p. diagrs. (NACA TM 1358. Trans. from Aerotecnica, v.31, no.4, August 15, 1951, p. 225230). The idea is advanced of making a supersonic nozzle by producing one, two, or three successive turns of the whole flow; with the result that the wall contour can be calculated exactly by means of the Prandtl Meyer "Lost Solution. ' 3 BRITISH REPORTS N20661* Aeronautical Research Council IGt. Brit.) NOTE ON PROFILE DRAG CALCULATIONS FOR LOWDRAG WINGS WITH CUSPED TRAILING EDGES. R. C. Lock 1952. 10p. diagrs., tab. (ARC R & M 2419; ARC 9772. Formerly RAE Aero 2130) In R. M. 1833 calculations of profile drag were made based on wing sections of conventional design, and were later extended in an Addendum to "low drag" wing sections with convex trailing edges. Further calculations were required for lowdrag sections of more recent design with cusped trailing edges. This report presents calculations made on sections of the NACA 65family of thickness 0. 12c and 0 23c with maximum thickness at 0. 4c from the leading edge, over a range of Reynolds number and position of the transition points. N20662' Aeronautical Research Council lGt Brit.) MODEL TESTS WITH FLOW ON THE GLOSTER F.9 40 WITH H. 1 NACELLES (METEOR H). J. S. Thompson. C. M. Fougere and E. G. Barnes. 1952. 17p. diagrs., 14 tabs. (ARC R M 2517; ARC 6901. Formerly RAE Aero 1821) For an earlier estimate by Thompson and Barnes of the longitudinal stability of the F. 9 40 in flight, it was necessary to extrapolate for the higher values of CL, because the maximum jet flow from model tests then available was that appropriate to a CL of only 0. 2, ground level. The purpose of the tests described in this note was to extend the previous model tests, using a considerably larger flow, to enable more precise estimates to be made. Meas urements were made of lift, drag, pitching and yaw ing moments for jet flows up to an exit v V of 4. 5 and at various tunnel speeds. N20664" Aeronautical Research Council (Gt. Brit.) THE DESIGN AND INSTALLATION OF SMALL COMA PRESSED AIR TURBINES FOR TESTING POWERED DYNAMIC MODELS IN THE ROYAL AIRCRAFT ESTABLISHMENT SEAPLANE TANK. D. I. T. P. LlewelynDavies, W. D, Tye and D. C. MacPhail. 1952. 19p. diagrs.. photos., tab. (ARC R & M 2620; ARC 10.812 Formerly RAE Aero 2192) This report describes the development of small light weight air turbines for powering dynamic models in the R. A. E. seaplane tank. The units have proved to. be rugged and reliable and power., weight ratios of 0.4 lb bhp have been achieved. The installation of the turbines in dynamic models and the provision of their air supply are also discussed. 4 N20665' Aeronautical Research Council (Gt. Brit.) ON THE SOLUTION OF LINEAR SIMULTANEOUS DIFFERENTIAL EQUATIONS WITH CONSTANT COEFFICIENTS BY A PROCESS OF ISOLATION. J. Morris. 1952. 7p. (ARC R & M 2623; ARC 11,420. Formerly RAE SME 4036) In this report, a process is given for the solution of linear differential equations with constant coeffi cients. The operative artifice is closely akin to Routh's method of isolation by means of which the constants of integration are found separately for each root of the characteristic equations. N20666" Aeronautical Research Council (Gt. Brit.) THE DYNAMIC LANDING LOADS OF FLYING BOATS WITH SPECIAL REFERENCE TO MEASURE MENTS MADE ON SUNDERLAND TX. 293. Anne Burns and A. J. Fairclough. 1952. 38p. diagrs., photos., 2 tabs. (ARC R M 2629; ARC 11,344. Formerly RAE Structures 17) An account is given of a fullscale investigation into the stresses occurring in the wing members of a Sunderland flying boat during landing impacts. It is found that the main dynamic effect is caused by the wing oscillating in its fundamental mode. These dynamic loads have a spanwise distribution similar to the normal lift load and, if the level flight lift load is taken as unity, a magnitude (in the most severe impact recorded) of 1.4 upwards and 1.5 downwards. Generalizing this result, one concludes that whereas down loads in landing may be a deciding factor in design the up loads are amply covered by existing requirements. Comparison of calculated and exper mental loads found in these tests indicates that sat isfactory agreement can be attained by using recent ly introduced modifications of standard dynamical methods. Although the investigation is primarily a structural one some interesting results on general water load phenomena are obtained. N20667* Aeronautical Research Council (Gt. Brit. LOAD DIFFUSION AT AN INTERSPAR OPENING: THEORETICAL METHODS OF ANALYSIS COM PARED WITH STRAIN MEASUREMENTS ON A LARGE WING. D. C. Allen. 1952. 26p diagrs., 6 tabs. (ARC R r M 2664; ARC 11,731. Formerly RAE Structures 30) The diffusion of load from spar flanges into skin and stringers near an opening was investigated experi .mentally in a large wing structure undergoing strength tests. A comparison of measured strains with those given by theoretical methods shows that in general the flange loads are represented with rea sonable accuracy. Any theory, however, in which the chordwise rib at the edge of the opening is ig nored gives shear stresses much greater than those measured. Allowance for the bending stiffness of this rib produces values of shear stress comparable with those obtained experimentally. NACA RESEARCH ABSTRACTS NO.8 N20668* 1 ,: : Aeronautical Research Council (Gt. Brit.) ". LANDING GEAR WITH TWIN TANDEM WHEq]L::. . UNITS: CORNERING CHARACTERISTICS AS DE: A TERMINED BY MODEL TESTS. J. W..Blii horn... 1952. 7p. diagrs., photos., tab. (ARC R&. M2668; ARC 11,850. Formerly RAE Tech. Note Meeh. Eng. 18) For twin tandem units, the wheel loading conditions which arise when aircraft are turned on the ground : i' may be critical for the landing gear. To estimate .:, the magnitude of these loads, cornering tests wer ... made on a smallscale model of the main under. carriage unit proposed for the Brabazon 1, MK. II. . These tests showed that for zero turning radius, that is, turning about the central vertical axis of the model undercarriage, the wheel side loads were al most equal to the vertical load multiplied by the ' coefficient of sliding friction between the tires and the ground. The side loads rapidly decreased as the . turning radius increased, and with the turning radius" "' equal to three times the wheel base, the wheel side loads were only about half of those at zero turning radius. The severity of the design loads for turning on the ground will therefore be considerably reduced .: if it can be ensured that the center of thd minimum turning circle of the aircraft is a short distahe ". outboard of either main undercarriage unit. N20669 Aeronautical Research Council (Gt. Brit.) CONICAL FLOW AS A RESULT OF SHOCK AND ', BOUNDARYLAYER INTERACTION ON A PROBE. J. Lukasiewicz. 1952. 16p. diagrs., photos.. (ARC R & M 2669; ARC 12, 023. Formerly RAE .. Tech. Note Aero 1968; SD 85) . The formation of a conical shock and a conical region.. of flow separation originating from the tip of a thin traversing tube was observed in a supersonic tunnel as a result of interaction of a strong shock with the. boundary layer on the tube surface. The angles of . the conical shock and separation surfaces and the static pressure in the separation region are in good . agreement with the theoretical conical flow solutions.. .' The extent of the conical flow illustrated should act . as a warning against the use of static pressurs tube'i;.: for measuring pressures in the regions of strong . shocks. ;. N20670' A ./.. .* Aeronautical Research Council (Gt. Brit.) AN ELECTRIC TANK FOR THE DETERMINATION OF THEORETICAL VELOCITY DISTRIBUTION T. J. Hargest. 1952. 9p. diagrs., photos. .A kC .:; R& M 2699; ARC 12,448. Formerly NOTE Miemo.: M. 48) ::.. ....i An analogy due to Relf has been applied to the design :... of apparatus for quickly determining the theoretici ::. velocity distributions around an airfoil in cascade.' The accuracy of the apparatus was tested by deteri"'^. mining the velocity distribution around a cylinder:' t' '," .. ..: : . NACA RESEARCH ABSTRACTS NO.38 An accuracy of within 1 percent of the approach velocity was obtained for this case. The apparatus .: has since been applied to determine the theoretical : velocity distribution around various airfoils in cas r cades; an example is given of the pressure distribu tion around an airfoil at zero incidence. An applica tion to determine the theoretical velocity distribution around the central airfoil of a nozzle cascade where the effect of the ducting side walls is included is also given. N20671*1 Aeronautical Researoh Council (Gt. Brit.) VELOCITY DISTRIBUTION ON STRAIGHT AND SWEPTBACK WINGS OF SMALL THICKNESS AND INFINITE ASPECT RATIO AT ZERO INCIDENCE. S. Newmark. 1952. 40p. diagrs. (ARC R M 2713; ARC 10, 907. Formerly RAE Aero 2200) A solution by H. Ludwieg, giving the velocity distri bution in the central section of a thin sweptback wing of infinite aspect ratio with a biconvex profile at zero incidence, has been found erroneous. In connection with this problem, the approximate method of sources and sinks for determining velocity distribu tion on straight and sweptback wings is critically examined, its limitations established, and proper ways of its application to threedimensional problems indicated. A correct solution of Ludwieg's problem is found, and generalized to give the velocity distri bution over the entire wing. The method is further extended to cover a wide class of thin symmetrical wing profiles, those with rounded leading edge being, however, often intractable by this particular method. The ultimate purpose of the investigation is to pro vide a reliable basis for determining the critical Mach number for sweptback wings. Further work is needed to embrace wings of finite aspect ratio and tapered wings, in particular, delta wings The method seems adequate to deal with these more com plex cases. N20672 * Aeronautical Research Council (Gt. Brit THE NUMERICAL SOLUTION OF TWO DIMENSIONAL FLUID MOTION [N THE NEIGH BOURHOOD OF STAGNATION POINTS AND SHARP CORNERS. L. C. Woods. 1952. 15p. diagrs. (ARC R & M 2726. Formerly ARC 12, 887, FM 1407) Methods are given in this paper of dealing with sin gularities of functions satisfying certain two dimensional partial differential equations. For a numerical solution, the differential equations are replaced by difference equations on a square mesh. Log (1/q) where q is the velocity, becomes infinite at stagnation points, sharp corners, sinks, etc., while the conjugate function a (flow direction) becomes multivalued. The method consists in finding a series expansion for the function (log 1. q or 0) in the neighborhood of the singularity. This expansion is then used to find relationships between the function values at points of the mesh adjacent to the singulari ty. A method of working directly in the transformed .5 flow plane tin which the airfoil is a slit), and thus avoiding irregular squares on the boundary, is also given. The method is developed for incompressible flow, but an approximation suitable for compressible flow is given. DECLASSIFIED NACA REPORTS NACA RM A8E17 THE ASYMMETRIC ADJUSTABLE SUPERSONIC NOZZLE FOR WINDT UNNEL APPLICATION. H. Julian Allen. July 23. 1948. 42p. diagrs, photos., 2 tabs. (NACA RM A8E171 (Declassified from Restricted, 10 30 52) An asymmetric adjustable nozzle ior supersonic wind tunnel application which permits continuous adjust ment of the testsection Mach number is described. The characteristics of this nozzle are compared with the more conventional supersonic tunnel nozzles. NACA RM L8E06 AERODYNAMIC CHARACTERISTICS AT HIGH SPEEDS OF RELATED FULLSCALE PROPELLERS HAVING DIFFERENT BLADESECTION CAMBERS. Julian D. Maynard and Leland B. Salters, Jr. August 31, 1948. 54p. diagrs., tab., photo (NACA RM L8E06) (Declassified irom Restricted, 11 26 52) Comparisons are made of results obtained in wind tunnel tests of related fullscale propellers over a range of blade angles from 200 to 550 at airspeeds up to 500 miles per hour to evaluate the combined effects of bladesection camber and compressibility on propeller aerodynamic characteristics NACA RM L8E07 AERODYNAMIC CHARACTERISTICS AT HIGH SPEEDS OF FULLSCALE PROPELLERS HAVING CLARK Y BLADE SECTIONS. Peter J. Johnson. October 26, 1948. 60p. diagrs photos., tab (NACA RM L8E07) (Declassified from Restricted, II 26 52) Results obtained in windtunnel tests of two full scale propellers over a range oi blade angles from 200 to 550 and at airspeeds varying from 60 to 485 miles per hour are presented NACA RM L8E24 AERODYNAMIC CHARACTERISTICS OF A TWO BLADE NACA 10(3)(08)03R PROPELLER. Albert J. Evans and Leland B. Salters, Jr. September 2, 1948. 29p. diagrs., tab. (NACA RM L8E24) (Declassified from Restricted, 11 26.'52) Contains results of windtunnel tests on a fullscale NACA 10(3)(08)03R twoblade propeller The w war ar UNIVERSITY OF FLORIDA 3 1262 08153 270 tests were part of a program to determine the effects of bladeshank design on propeller aerodynamic characteristics. A maximum efficiency of 91.5 per cent was attained at a rotational speed of 1600 revolutions per minute at a 300 blade angle. Peak efficiency at a blade angle of 450 was decreased 32 percent by increasing the hclicaltip Mach number from 0.80 to 1.20. NACA RM L8H16 AERODYNAMIC CHARACTERISTICS OF A THREE BLADE PROPELLER HAVING NACA 10(3)(08)03 BLADES. Robert E. Davidson. October 29, 1948. 29p. diagrs., tab. (NACA RM L8H16) (Declassified from Confidential, 11 26, 52) Contains results of windtunnel tests of a 10foot diameter, threeblade propeller at stream Mach numbers from 0.12 to 0.64. The tests were part of a program to determine the effects of blade number on propeller aerodynamic characteristics. The pro peller blades are of designation NACA 10(3)(08)03. A maximum efficiency of 92 percent was attained at a rotational speed of 1350 rpm at a 400 blade angle. Peak efficiency at the design blade angle of 450 was decreased 4 percent by increasing the helical tip Mach number from 0.80 to 0.96. NACA RM L9G20 TWODIMENSIONAL WINDTUNNEL INVESTIGA TION OF A 6PERCENTTHICK SYMMETRICAL CIRCULARARC AIRFOIL SECTION WITH LEADINGEDGE AND TRAILINGEDGE HIGH LIFT DEVICES DEFLECTED IN COMBINATION. Robert J. Nuber and Gail A. Cheesman. September 6, 1949. 29p. diagrs., photo.. 3 tabs. (NACA RM L9G20) (Declassified from Restricted, 10'7/52) An investigation ol a 6percentthick symmetrical circulararc airfoil with leadingedge and traillng edge highlift devices was made to determine the effectiveness of these devices in increasing the maximum section lift coefficient of the airfoil when deflected in combination. The results indicated that, with the plain trailingedge flap deflected 600, maxi mum section lift coefficients of 2.02 and 1.95 can be obtained by deflecting a 15percentchord leading edge slat or a 15percentchord droopednose flap, respectively. Variations in Reynolds number for either the slat or droopednoseflap configurations or moving the droopednoseflap hinge from the lower surface to the upper surface had essentially no effect on the lift characteristics. NACA ..* ' RESEARCH ABSTRACTS NO.="f;..l : : :""' ' .. '" '" ; ... v..... :T. i',, . AA r.. ., . :. : .., : i" .. ':. .: :.. ,.. .7 : ,. . .. :: .. ..J i .'. *. ...... '" I S .' .' S t :.,.* .: *":.'.:::,:** NACAL4 . : *11 5 a i3 : ::. : ., ..... ....t~w .,n . 
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