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Aviation History
1955
1955 - 0844.PDF
842 FLIGHT, 17 June 1955 SINGLE-SPOOL TURBOPROP DECREASING STAGE EFFICIENCY x 8 9 10 II DESIGN PRESSURE RATIO -;. DESIGN POINT,. COMPRESSOR OPERATING LINE COMPRESSOR SuRGE LINE 100 h 90°U 80% ADIABATIC EFFICIENCY 70°/. SPEED •''--' -..:.". . OESIGN COMPRESSOR OPERATING LINE COMPRESSOR SURCE-^/, LINE yu POINT \ 70°/ //S\ j'mx A\9O% 80% \ ADIABATIC EFFICIENCY SPEED MASS FLOW MASS FLOW Figs. 1, 2, 3 and 4 (clockwise): Effect of pressure ratio upon specific fuel consumption for a typical modern turboprop; characteristics of a compressor in which the optimum efficiencies are close to the surge; a compressor characteristic in which optimum efficiencies are well away from the surge; build-up of fuel flow during acceleration for engines with various ••• compressor characteristics. this was overcome by maintaining constant r.p.m. over the normaloperating range. This had the disadvantage that for low-output conditions the economy was poor, and that the airscrew had ahigh tip speed (and high noise level) during ground idling. Further, such an engine would always be prone to surge at someaircraft-operating condition, such as a particular flight attitude, where the compressor characteristics might be influenced by theintake conditions. In order to take full advantage of the "Fig. 3 type" compressorit was essential to employ a comprehensive fuel control system. Another factor was the importance of uniform temperature dis-tribution by the exit from the combustion chamber and it was also dssifsble to minimize combustion pressure losses—although thelatter should not be achieved to the detriment of satisfactory relighting at high altitude. Potentially there was much more performance development ina new turboprop than in the piston engine, and experience had shown that this could be achieved in a shorter -time. Increasedambient temperatures reduced turboprop performance more than that of the piston engine, but water-methanol could restore lostpower. By employing internally cooled turbine blading, engines could be operated with existing materials at higher turbine-inlettemperatures and an increase in output of up to 50 per cent could be achieved, with improved economy. There was an optimumratio of jet thrust to shaft power for any given flight speed; for example, the figures for cruising at 300, 400 and 500 km/hr were7, 9 and 12 per cent respectively. The jet-thrust percentage could be adjusted by varying the area of the propelling nozzle;unfortunately, the run of the jet-pipe, in some installations, might preclude the attainment of the optimum jet thrust, and it wasapparent that tail-pipe design became increasingly important at higher cruising speeds and altitudes. A major requirement of the fuel control system was that thepilot should have full control of power from a single lever, with all safeguards and changes due to operational conditions beingeffected automatically. The fuel metering and airscrew control systems should be suitably interconnected with safeguards againstsurging and /or over-heating during automatic starts, rapid accelerations and normal operation; safeguards should also beprovided against over-speeding and excessive torque loading. It was essential, during development of the control system and Fig. 5. Napier Eland turboprop, showing easy break-down into principal assemblies. CURVE A NORMAL OPERATION CURVE B ACCELERATION FOR COMPRESSOR OF FIG. 5 CURVE C ACCELERATION FORCOMPRESS08 OF FIG. 2 CURVED MAXIMUM TEMPERATURE LIMITATION IPUHG ENfiINt ft.P.rt. general handling, tohave a full know- ledge of the actualcharacteristics of the main componentsof the engine. Turning next tomechanical design, Mr. Penn pointedout that, compared with the pistonengine, the turbo- prop had few cem=ponents subject to wear, the major problems being those of metal fatigue and creepat high temperatures. Owing to the much higher rotational speeds, a more accurate control had to be maintained on theconcentricity and timing balance of all rotating components. (For example, an out-of-balance load of the order of 53 lb wouldbe imposed on each bearing of a compressor shaft of 3.8in external diameter, 3.2in bore and 19iin long, which had a0.0008in eccentricity in the bore.) Consideration had also to be given to the relative movements between adjacent parts resultingfrom the large and rapidly changing temperature differences throughout the engine. Other design problems arose from theemployment of air and gas seals for rapidly revolving components and the whirling problems of the main rotating system, resultingfrom the fact that there was usually an appreciable span between bearing supports which, owing to the necessity of having alight structure, could not be completely rigid. Bearing cooling and oil sealing also demanded careful design. At this point, Mr. Penn turned his attention to his own com-pany's product, the Napier Eland. (This engine was fully described in our issue of July 23rd, 1954.) The design of thisengine was started at the end of 1950 with a design power of 3,000 e.h.p. The first Eland ran in August 1952, since whenover 3,000 hours' development running and a 150-hr rehearsal type test had been completed. Flight development was at presentbeing carried out with a Varsity and an Elizabethan, and CV-340 were in the process of being equipped with two Elandseach (Flight, January 28th, 1955) and would fly this summer. Consideration was being given to applyingthe Eland to other fixed-wing aircraft and a special development for the Fairey Roto-dyne helicopter (Flight, December 24th, 1954) would be installed in the prototype bythe end of this year. There was also a 4,000 e.h.p. version of the Eland under develop-ment, and this engine would have an air- cooled turbine and improved fuel economy. At the front of the standard Eland was areduction gear with epicyclic gearing incor- porating a unique design of hydraulic torque-meter. The main rotating assembly consisted of a ten-stage compressor giving a 7:1 pres-sure ratio with a mass flow of 31 lb/sec and a three-stage reaction turbine designed tooperate at 1,165 deg K gas temperature and with an expansion ratio of 5.5:1. Six separatecombustion chambers were employed, of the Lucas upstrearn-injection type. With con-siderations of maintenance and servicing in
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