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Aviation History
1933
1933 - 0661.PDF
SEPTEMBER 28, 1933 «.--. THE AIRCRAFT ENGINEER SUPPLMIXNT TO FLIGHT TECHNICAL LITERATURE SUMMARIES OF AERONAUTICAL RESEARCH COMMITTEE REPORTS These Reports are published by His Majesty's Stationery Office, London, and may be purchased directly from H.M. Stationery Office at the following addresses: Adastral House, Kingsway, W.C.2; 120, George Street, Edinburgh; York Street, Manchester; 1, St. Andrew's Crescent, Cardiff; 15, Donegall Square West, Belfast; or through any Bookseller. OIL COOLING FOE AIRCRAFT. By B. C. Carter, F.R.Ae.S., M.I.Mech.E. Communicated by the Director of Scientific Research, Air Ministry. R. & M. No. 1486. (58 pages and 36 diagrams.) September, 1931. Price 3s. 6d. net. This report has been prepared as a review of means of oil cooling for aircraft and as an analysis of the problem of oil cooling. It is comprehensive in its survey and deals with all the main problems on this subject as affecting present day engines. For convenience the subject has been grouped into three sections. In Section 1, the chief basic quantities involved are reviewed and some values for cooling coefficients, etc., are given for reference purposes. In Section 2, a quantitative analysis is made of the rate of transference of heat from oil to metal and from metal to air. This analysis is based on simple theory, using the results of certain tests. A form of graph has been devised for depicting the characteristics of oil coolers as regards heat trans ference. Some estimates are made of thickness of virtual laminar layers of oil, in contact with the metal surfaces of coolers, through whieh the heat passes by conduction only. This matter is considered in fuller detail in Appendix I, and the transference of heat from a fluid in laminar motion is examined in Appendix II. In Section 3, types of oil coolers in use, under trial, or proposed, are described. Some information is gfven concerning the performance of these coolers, and in certain cases the results of tests are analysed. In Appendix III, the use of an intermediate fluid in the cooling of oil is considered and some test results are given. Little difficulty oresents itself in arranging to dissipate two or three horse power from the oil, but for higher rates of dissipation, particularly in a warm climate, an oil cooler tends to become bulky and heavy, and possibly to have excessive drag. The importance of drag depends upon the speed of the machine, as well as upon the design of the cooler, because the power absorbed in drag increases almost as the cube of the airspeed, whereas the heat dissipation increases at a lower rate than the first power of the airspeed. Put otherwise, the ratio of the power absorbed in drag to that represented by the rate of heat dissipa tion, increases practically as the square of the speed. At low aircraft speeds, this ratio is so small that the power lost in drag is negligible, but, at higher speeds, the drag horse-power may be very important if the cooling surface is not part of the aircraft surface. For racing aircraft, it is essential to employ part of the aircraft surface for nil cooling, and for high speed non-racing aircraft, it is desirable to do the same. To get the beet results involves making arrangements for oil cooling at an early stage of the machine design, with sufficient information available tn enable the cooler performance to be predicted. At present there are important gaps in our knowledge of the subject, and a corresponding element of doubt concerning the provision that needs to be made for oil cooling. The amount of horse-power that will require to be dissipated from the oil to prevent excessive feed temperatures is largely a matter of surmise because it depends upon the degree of cooling of the engine as installed. A considerable amount of cooling is often effected in the oil tank and pipes. Thus, when oil temperature measurements are made, the location of the thermometers has an important bearing on the interpretation of the results. Kven if the horse-power to be dissipated is known, the size of cooler required cannot be estimated within very close limits, except from previous practice, because the metal-air and oil-metal heat transfer coefficients depend upon many factors. From the values of these coefficients given in this report, it should be possible to make some progress towards superseding trial and error methods and towards the adoption of oil coolers which do not involve added drag. Where oil is sprayed on a surface above the oil level in the tank and drains away from the surface, the cooling element cannot become filled with congealed oil. Thus the danger is eliminated of the cooler ceasing to control the feed oil temperature under all flight conditions. The method has a further advantage which concerns vulnerability. In the event of the cooling element being punctured, the rate of oil leakage is small in relation to the rate at which the total oil content of the system is circulated. When the sprayed surface is below the oil level in the service tank, no air vent is provided and the cooler acts as an air vessel under the small pressure needed to force the cooled oil through a pipe discharging above the oil in the tank. Spraying introduces its own difficulties and these should be mentioned. To get the best results from spraying involves using a pressure of 30 to 50 lb. per sq. in., and where scavenge pumps have not been developed to operate against a back pressure of this magnitude, troubles have arisen in isolated instances with frothing, cavitation at the scavenge pump inlet and pressure fluctuation. These have been due to the presence of air and are peculiar to the engines concerned. Many other tests have been made without such troubles occurring. They should not arise if a separate pump, running full, were fitted for circulating the oil through the spray cooler (and cleaner where incorporated in the cooler) and back to the engine sump. With this arrange ment a high rate of circulation may be adopted to obtain vigorous spraying when the rate of heat dissipation necessitates spraying a large area. If an oil cooler under trial is found not to give the amount of cooling required it is desirable to determine the coefficients of heat transference from oil to metal and from metal to air. The latter coefficient may be found in wind tunnel tests by circulating water vigorously instead of oil, and such Jests incidentally give the upper limit to the possible performance with oil. The oil-metal coefficient for each wind tunnel test made with oil may then be deduced. An analysis made in this way should indicate any changes that would give improved performance for the conditions under which the cooler i« to be used. It is for consideration whether further wind tunnel tests be made on the more important types of cooler that project into the relative wind, in order to determine the drag, and the coefficients of heat transference under different conditions as regards wind speed, temperatures and rates of oil circulation. The oil-metal coefficient for various oil and metal temperatures needs to be correlated more closely with the Reynolds number and other non-dimensional quantities for specific types of oil coolers, and further experiments might be made with this objective. These would not involve the use of a wind tunnel. In preparing this report, the author has drawn freely upon the result.- o tests made by his colleagues at the Royal Aircraft Establishment in the Departments concerned. EXPERIMENTS ON SWEPT-BACK AND SWEPT-FORWARU AEROFOILS. By D. H. Williams, B.Sc., and A. S. Halliday, B.Sc., Ph.D., D.I.C. With an Appendix bv B. B. Irving, B.Sc. R. & M. No. 1491. (22 pages and 21 diagrams.) October, 1930. Price Is. 3d. net. In the Appendix to this report, H. B. Irving has given some results deduced from pressure plotting data on the effects of sweep-back and sweep-forward on the aerodynamic characteristics of an aerofoil, showing that a swept - forward wing might possess marked advantages over a straight wing or one with sweep-back. A delayed stall and an increase in lift at high angles of incidence due to sweep-forward were indicated and also an increase in the rolling moment due to sideslip above the stall. The experiments described below were carried out to test the conclusions by determining the rolling moments due to sideslip and roll on an aerofoil with various angles of sweep- back and sweep-forward. The wings tested were rectangular, aspect ratio 0, with square wing tips. They were mounted in two parts on a metal plate in such a way that each hall' could be rotated independently about an axis normal to the aerofoil in the centre section up to 30° each way. In this way, wings with any angle of sweep-back or sweep-forward could be constructed, but each wing tip section was at right angles to the centre line of the corresponding half of the wing ; the wing tip section was, therefore, not along the wind direction at 0° yaw, except for the straight wing. The general conclusions deduced from the experiments may be summarised us follows:— (1) Sweep-back and sweep-forward decrease the lift below the stall and the maximum lift coefficient, but they increase the lift above the stall. (2) 10° sweep-back halves the value of Ip above the stall but has little effect below the stall. 10° sweep-forward increases Iv above tiie stall and reduces it below the stall. (3) Moderate sweep-back tends towards instability in roll, while sweep- forward postpones the change from stability to instability. A STUDY OF AIRCRAFT TURNING PERFORMANCE. PART I. By S. B. Gates, M.A. Communicated by the Director of Scientific Research, Air Ministry. R. & M. No. 1502. (8 pages and 11 diagrams.) August 10, 1932. Price 6d. net. A method of calculation is proposed by means of which the steady spiral motion of an aircraft at small angles to the horizon can be simply deduced from knowledge of its angles of climb and glide in rectilinear flight. Curves are shown illustrating (a) true banked turns at various rates of descent, and (ft) level turns with various degrees of sideslip, A rough estimate ia made Of thfi effect Of wing-tip slots. Equilibrium of forces only is considered, and the limitations introduced by available control will be Studied in a later report. PERFORMANCE TESTS OF CERTAIN EXPERIMENTAL DESIGNS OF DIFFUSER AND IMPELLER IN A CENTRIFUGAL SUPERCHARGER, WITH PARTICULAR REFERENCE TO THEIR INFLUENCE UPON SURGING, INCLUDING THE EFFECT OF AN IMPRESSED PERIODICITY OF FLOW. By G. V. Brooke, B.Sc.Tech. Communicated by the Director of Scientific Research, Air Ministry. R. & M. No. 1503. (43 pages and 28 diagrams.) December, 1932. Price 2s. 6d. net. The majority of centrifugal superchargers constructed for aero-engine service have utilised diffusers containing either a comparatively large number of short, straight vanes or a smaller number of curved vanes of greater length. In bench calibration tests of such superchargers at constant rotational speed of the impeller, in which the mass flow of air is progressively reduced by restricting the outlet from the supercharger, it is found that the pressure of the air at delivery increases as the mass flow diminishes until a point is reached at which the discharge pressure decreases abruptly. This sudden breakdown in the character of the relation between delivery pressure and mass flow is generally termed the " surge point." As a rule it is accompanied by an audible air vibration, the severity of which is greatly Influenced by the volume oi air enclosed in the pipe system between the supercharger and the valve used for throttling the air flow. Testa were undertaken to determine the comparative performance of a supercharger and the relative proximity of the air flows corresponding respectively to maximum performance and to the inception of surging when several experimental types of diffuser and impeller were incorporated. The components tested included both shrouded and unshrouded impellers having either curved or straight radial blades, and diffusers (1) of simple vaneless type; (2) containing straight vanes set at various angles ; and (3) constructed in the form of two equiangular spiral volutes. The simple vaneless diffuser was beneficial in regard to the postponement of surging, but produced very low values of adiabatic efficiency and pressure ratio of compression. A diffuser of this type is unsuitable for use in a high speed compressor of small overall diameter, as the length of diffuser passage available is insufficient for reduction of the final air velocity to the required value. The performance of the supercharger was influenced to a considerable extent by change of the difluser vane angle. Improvements in compression ratio and efficiency were obtained from the shrouded impellers. In particular, the design incorporating curved blades afforded a very high compression ratio, but its maximum value at each impeller speed was reached at an air flow in close proximity to that at which surging commenced. Comparing the performanees of the supercharger on the basis of the useful power which it enables a hypothetical engine to develop, and excluding the cases of the plain vaneless diffuser and the less favourable vane angles, no 972 e 1
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