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Aviation History
1928
1928 - 0447.PDF
MAY 31, 1928 43 THE AIRCRAFT ENGINEER SUPPLEMENT TO FLIGHT depends upon the viscosity of the oil upon the cylinder walls, which in turndepends upon the temperature of the jacket water. (6) While theoretical considerations would lead one to expect an increase in friction with increasein compression ratio the evidence at hand incicates that this effect is slight The second section of the report deals with measurements of the frictionof a group of pistons differing from each other in a single respect such as length, clearness, area of thrust face, location of thrust face, etc Resultsobtained with each type of piston are discussed and attention is directed particularly to the fact that the friction chargeable to piston rings dependsupon piston design as well as upon ring design. This is attributed to the effect of the rings upon the thickness and distribution of the oil film which,in turn, affects the friction of the piston to an extent which depends upon its design. T.R. No. 263. " PRELIMINARY FLIGHT TESTS OF THE N.A.C.A. ROOTS TYPE AIRCRAFT ENGINE SUPERCHARGER." By Arthur W. Gardiner and Elliott G. Reid, N.A.C.A. An investigation of the suitability of the N.A.C.A. Boots type aircraftengine supercharger to flight-operating conditions, as determined by the effects of the use of the supercharger upon engine operation and aeroplaneperformance, is described in this report. The supercharger has been previously described in N.A.C.A. Technical•Report No. 230 ; the results of laboratory tests are also given there. The compressor has a displacement of 0-51 cub. ft. per revolution, and weighs88 lbs. The selection of a suitable propeller and the provision of satisfactoryintake ducts and adequate engine cooling were preliminary problems. The supercharger was first tested in a modified DH 4 aeroplane'with a 5-4 com-pression-ratio " Liberty-12 " engine. Two sets of drive gears, which enabled the maintenance of sea-level pressure at the carburettor intake up to12,000 and 20.000 ft., were provided. The higher gear ratio supercharger was next tested in a DT.2 aeroplane, which was later converted into atwin-float seaplane ; the DT.2 also had a " Liberty " engine. Loads up to 2,000 lbs. were carried in the seaplane with normal and superchargedengines. Attention was concentrated on the operation of the engine-superchargerTinit, and on the improvement of climbing ability; some information concerning high speeds at altitude was obtained.The supercharger was found to be satisfactory under flight-operating conditions. Although two failures occurred during the tests, the causesof both were minor and have been eliminated. Careful examination of the engines revealed no detrimental effects which could be attributed to super-charging. Marked improvements in climbing ability and high speeds at altitudewere effected. It was also found that the load which could be carried to a given moderate or high altitude in a fixed time was considerably augmented.A slight sacrifice of low-altitude performance was necessitated, however, by the use of a fixed pitch propeller. From a consideration of the very satisfactory flight performance of theRoots sui>ereharger and of its inherent advantages, it is concluded that this type is particularly attractive for use in certain classes of commercialaeroplanes and in a number of military types. T.R. No. 264. " DIFFERENTIAL PRESSURES ON A PITOT- VENTURI AND A PITOT-STATIC NOZZLE OVER 360° PITCH AND YAW." By R. M. Bear, Construction Department, Washington Navy Yard. Measurements of the differential pressures on two Navy air-speed nozzles,consisting of a Zahm-type Pitot-Venturi tube and a SQ.16 two-pronged Pitot-static tube, in a tunnel air stream of fixed speed at various angles ofpitch and yaw between 0" and ± 180°, show for a range over — 20° to + 20° pitch and yaw, indicated air speeds varying very slightly over2 per cent, for the Zahm type and a maximum of about 5 per cent, for the 3Q.16 type from the calibrated speed at 0°. For both types of air-speed nozzle the indicated air speed increasesslightly as thetubes are pitched or yawed several degrees from their normal 0° attitude, attains a maximum around ± 15° to 25°, declines rapidlytherefrom as ± 40° is passed, to zero in the vicinity of ± 70° to 100°, and thence fluctuates irregularly from thereabouts to ± 180°. The completevariation in indicated air speed for the two tubes over 360° pitch and yaw is graphically portrayed in Figs. (1 and 10. For the same air speed and 0° pitch and yaw the differential pressureof the Zahm type Pitot-Venturi nozzle is about seven times that of the SQ.16 type two'-pronged Pitot-static nozzle. T.R. No. 265. " A FULL-SCALE INVESTIGATION OF GROUND EFFECT." By Elliott G. Reid, N.A.C.A. This report describes flight tests which were made with a Vought VE.7aeroplane to determine the effects of flying close to the ground. It is found that the drag of an aeroplane is materially reduced uponapproaching the ground, and that the reduction may be satisfactorily calculated according to theoretical formulas. Several aspects of ground effect which have had much discussion are explained. T.R, No. 266. " AIR FORCE AND MOMENT FOR N-20 WING WITH CERTAIN CUT-OUTS." By R, H. Smith, Construction Department, Washington Navy Yard. The aeroplane designer often finds it necessary, in meeting the requirementsof visibility, to remove area or otherwise locally to distort the plan or section of an aeroplane wing. This report, prepared for the Bureau of Aeronautics.January 15, 1925. contains the experimental results of tests on six 5 by 30 in. N.20 wing models, cut out or distorted m different ways, which wereconducted in the 8 ft. by 8 ft wind tunnel of the >.avy Aerodynamical Laboratory in Washington in 1924. ., The measured and derived results arc given without correction for \i_« or for wall effect and for standard air density p = 0-00237 slug per cub. ft. T.R. No. 267. "DRAG OF WINGS WITH END PLATES." By Paul E. Hemke, N.A.C.A. In this report a formula for calculating the taduced drag of niultiplaneswith end plates is derived. The frictional drag of the end Plate8.^ al8? calculated approximately. It is shown that the reduction "f^e induceddrag, when end plates are used, is sufficiently large to increase the efflciencj of the wing. _ „ , .. *__ ~in^na an(j biplanes are - of end plate are determined for typical cases. The method of obtaining thereduction of drag for a multiplane is described. Comparisons are made of calculated and experimental results obtained inwind-tunnel tests with airfoils of various aspect ratios and end plates of various sizes. The agreement between calculated and experimental resultsis good. Analysis of the experimental results shows that the shape and sectionof the end plates are imjiortant. T.R. No. 2C8. " FACTORS IN THE DESIGN OF CENTRIFUGAL- TYPE INJECTION VALVES FOR OIL ENGINES." By W. F. Joachim and E. G. Beardsley, N.A.C.A. This research was undertaken at the Langley Memorial AeronauticalLaboratory, in connection with a general study of the application of the fuel injection engine to aircraft. The purpose of the investigation was todetermine the effect of four important factors in the design of a centrifugal type automatic injection valve on the penetration, general shape, and dis-tribution of oil sprays. The general method employed was to record the development of singlesprays by means of special high-speed photographic apparatus capable of taking 25 consecutive pictures of the moving spray at a rate of 4,000 persecond. Investigations were made concerning the effects on spray character- istics of the helix angle of helical grooves, the ratio of the cross-sectionalarea of the orifice to that of the groove?, the ratio of orifice length to diameter, and the position of the seat. The sprays were injected at 6,000, 8,000,and 10,000 lbs. per square inch pressure into air at atmospheric pressure and into nitrogen at 200, 400, and 600 lbs. per square inch pressure. Orificediameters from 0 -012 to 0 040 inch were investigated. It was found that decreasing the pitch of the helical grooves and thusincreasing the centrifugal force applied to the spray increased the spray cone angle considerably, although the percentage increase was much lessin dense air than in the atmosphere. On the other hand, the spray pene- tration decreased with increase in the amount of centrifugal force applied.About twice as much spray volume per unit oil volume was obtained with a high centrifugal spray as with a noncentriiugal spray. The spray coneangle increased, and the spray volume to oil volume ratio and spray pene- tration decreased with increase in the ratio of orifice area to groove area.Maximum spray penetration was obtained with a ratio of orifice length to diameter of about 1-5, Slightly greater penetration was obtained with theseat directly before the orifice. T.R. No. 269. " Am FORCE TESTS OF SPERRY MESSENGER MODEL WITH SIX SETS OF WINGS." BV James M. Shoe- maker, N.A.C.A. The purpose of this test was to compare six well-known aerofoils, the R.A.F.15, U.S.A. 5, U.S.A. 27, U.S.A. 35-B, Clark Y, and Gottingen 387, fitted to the Sperry Messenger model at full-scale Reynolds Number as obtained inthe variable-density wind tunnel of the National Advisory Committee for Aeronautics : and to determine the scale effect on the model equipped withall the details of the actual aeroplane. The results show a large decrease in minimum drag coefficient upon increasing the Reynolds Number fromabout one-twentieth scale to full scale. Maximum lift coefficient was in- creased with increasing scale for all the aerofoils except the Gottingen 3F7,for which it was slightly decreased. A comparison is made between the results of these tests and those obtained from tests made in this tunnel onaerofoils alone. T.R. No. 270. " THE MEASUREMENT OF PRESSURE THROUGH TUBES IN PRESSURE DISTRIBUTION TESTS." By Paul E. Hemke, N.A.C.A. The tests described in this report were made to determine the error causedby using small tubes to connect orifices on the surface of aircraft to central pressure capsules in making pressure-distribution tests.Aluminium tubes of three-sixteenths inch inside diameter were used to determine this error. Lengths from 20 ft. to 226 ft. and pressures whosemaxima varied from 2 ins. to 140 ins. of water were used. Single-pressure impulses for which the time of rise of pressure from zero to a maximumvaried from 0-25 sec. to 3 secB. were investigated. The results show that the pressure recorded at the capsule on the farend of the tube lags behind the pressure at the orifice end and experiences also a change in magnitude. For the values used in these tests the time lagand pressure change vary principally with the time of rise of pressure from zero to a maximum and the tube length. Curves are constructed showingthe time lag and pressure change. Empirical formulas are also given for computing the time lag. Analysis of pressure-distribution tests made on aeroplanes in flight showsthat the recorded pressures are slightly higher than the pressures at the orifice and that the time lag is negligible. The apparent increase in pres-sure is usually within the experimental error, but in th» case of the modem pursuit type of aeroplane the pressure increase may be 5 per cent. Forpressure-distribution tests on airships the analysis shows that the time lag and pressure change may be neglected. T.R. No. 271. "PRESSURE DISTRIBUTION TESTS ON PW-9 WTING MODELS SHOWING EFFECTS OF BIPLANE INTER- FERENCE." By A. J. Fairbanks, N.A.C.A. In this report tests are described in which the distribution of pressuresover models of the wings of the PW-fi aeroplane was investigated. The wing models were tested individually and in the biplane combination. The in-vestigation was conducted in the amtospheric wind tunnel of the National Advisory Committee for Aeronautics. It is concluded in this paper thatthe effect of biplane interference on the pressures on the wings is practically confined to the lower surface of the upper wing and the upper surface ofthe lower wing : that the overhanging portion of the upper wing is not greatly affected by the presence of the lower wing ; and that a slight washing atthe centre section of the upper wing satisfactorily compensates for a reduced chord at this section (providing the aerofoil section is not mutilated) andprevents a large reduction in the normal force over this portion of the wing. T.R. No. 272. " THE RELATIVE PERFORMANCE OBTAINED WITH SEVERAL METHODS OF CONTROL OF AN OVER-COMPRESSED ENGINE USING GASOLINE." By Arthur W. Gardiner and William E. Whedon, N.A.C.A. This Teport presents some results obtained at the Langley Memorial Aero-nautical Laboratory during an investigation to determine the relative 404g
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