Rolls-Royce Gas Turbines

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The first active interest in jet propulsion was shown by the Rolls-Royce organisation in 1938 when a department was established for the design of gas turbines. By 1940 test rigs for airfoils, bearings and combustion chambers had been set up and toward the end of the year the Company was manufacturing components for Whittle units. Next an intensive study of radial compressors for these units was undertaken to ascertain the causes of, and the means of eliminating, surging. A special test plant was installed, with a 2,000 h.p. Vulture engine driving the compressors. Late in 1941, under Air Ministry direction, the Company undertook to build a Whittle-type engine known as WRI, designed with low blade stresses to demonstrate that the gas turbine could be made completely reliable.
Early in 1943 Rolls-Royce took over research on the W2B/23 unit from the Rover Company whose engineers had developed straight-through combustion. Units of this type, with the name Welland (Rolls-Royce having decided to standardise river names for their gas-turbine nomenclature) were supplied for installation in the Gloster E.28. In June 1943 two units were fitted in the Gloster F9/40, prototype of the Meteor. This unit had a reverse-flow combustion system, a maximum diameter of 43 inches and could develop 1,700 Ib. thrust, although for the F9/40 it was de-rated to 1,450 Ib. By May 1944 the Welland was being regularly delivered to the R.A.F. Whilst all these activities were proceeding Rolls-Royce were engaged on a new design to utilise experience gained from development work. The new project was to be of the same maximum diameter for installation in the standard Meteor engine nacelles but to develop a static thrust of 2,000 Ib.
Drawings were commenced in April, 1943, and by July the unit was ready for test, and in November, 1943, it passed its 100-hour type-test at 2,000 Ib. thrust. In April of the next year it completed its first flight tests in the Meteor with a service rating of 1,800 Ib. thrust and a weight of 920Ib. The new engine was known as the B37, R-R Derwent, Series I. The Series II engine gave an increased thrust of 2,200 Ib. Series III was a special unit for experiments to provide suction on aircraft wing surfaces for boundary layer removal and Series IV gave a further increase in thrust to 2,400 Ib. The Derwent Series V, whilst retaining the maximum diameter of 43in., was an entirely new unit developing twice the thrust of the original Derwent I It is this unit which enabled the Gloster Meteor to achieve 606 m.p.h. (975 km/h.)

Derwent V

V15_files/derwent-perfomance.jpg This modern unit is, in effect, a scaled-down version of still another new type, the Nene. lts development was prompted by the promise shown by the Nene and the proof that the Meteor could utilise thrusts greatly in excess of the original estimates. A double-entry radial compressor of increased capacity as compared with previous Derwents, with an impeller about 21 in. diameter and twenty-nine radial blades on each side, is used on the Derwent V. At the other end of the shaft is a single-stage turbine. The main shaft is mounted in two roller bearings and a central ball-thrust bearing. Air is induced on both sides of the impeller and fed past the diffuser necks to the combustion chambers. Means are provided to cool the internal mechanism, including the centre and rear bearings and the front face of the turbine disc. A small centrifugal fan mounted in front of the centre bearing induces atmospheric air through short stub pipes on the front end of the engine housing and forces it through the cooling air manifold to the exhaust outlet at the rear. When it is realised that the turbine rotor runs in a gas temperature of about 850 deg. C., and that the 54 individual blades, measuring about 3in. long and 1.25in. wide, have to transmit about 75 h.p. the importanee of the metallurgical problems and the need for internal cooling will be understood. Apart from the anti-corrosive nature of the nickel chromium alloys used, non-creep properties are of the highest importance. A high-tensile strength must be maintained even under high working temperatures, as centrifugal force due to high speed of rotation imposes a heavy mechanical stress. Because of this high speed of rotation the compressor impeller and turbine rotor each needs to be statically and dynamically balanced, both individually and collectively as a single assembly. This explains the fact that the two shafts carrying compressor and turbine are connected by a quickly detachable toothed coupling. The impeller, which has the larger diameter, has a tip velocity of approximately 1,500 ft./sec. - that is considerably in excess of sonic speed. The installed weight of the Derwent V engine is under 1,500 Ib. and it delivers 4,000 Ib. thrust - a power/weight ratio never previously attained. For the world's speed record, two of these units in the Meteor developed sufficient power to attain 606 m.p.h. when throttled down to 3,600 Ib./thrust. Fuel consumption on the record was high, as full throttle low altitude conditions are the least favourable to thermal efficiency.

Nene X

Early in 1944 the Ministry of Aircraft Production issued a specification for a jet propulsion unit having a maximum overall diameter of 55in., a minimum static thrust of 4,000 Ib. and a weight not exceeding 2,200 Ib. The Rolls-Royce Nene I is the fulfilment of this requirement in generous measure. Units at present in production are 49½ in. diameter, develop a thrust of 5,000 Ib. and weigh only 1,550 Ib. Thus the realisable performance is 3.2 Ib. thrust per Ib. weight and 375Ib. thrust per square ft. of frontal area instead of 1.8 Ib. and 242 Ib. thrust respectively as originally stipulated. In the remarkably short period of 5½ months the design was completed, all drawings prepared, the first unit built and the proving run of one hour at 5,000 l]b. thrust successfully accomplished. The single stage, double-sided, radial flow compressor of the Nene delivers air at four times atmospheric pressure to nine straight-flow combustion chambers. Aviation kerosene under high pressure is sprayed downstream into the chambers, to form a combustible mixture having an air/fuel ratio of about 18: l, and burnt continuously. The major volume of the air, diluting the mixture to a ratio of approximately 60: l is expanded by the heat released by combustion of the fuel. The complete rotating assembly comprises the compressor impeller, cooling fan, and turbine rotor on two coupled shafts supported in three bearings. End bearings are of the roller type whilst the centre one is a deep groove ball bearing to support axial loads. It is of interest to note that at speeds up to about 8,000 r.p.m. the axial thrust is directed forward, but above that figure it is exerted rearward.
The impeller is 28.5in. diameter and machined with 29 radial vanes each side from a single light alloy forging. Curved entry vanes, approximately 17.75in. diameter, for each side are separate components machined all over. On the 13in. diameter cooling air fan the 30 vanes have integral entry sections which are bent cold in two stages with an intermediate annealing operation.
The compressor casing is built up of front and rear members attached to a central diffuser ring by bolts passing through the diffuser vanes and the intermediate sputter vanes. To facings at the nine outlets from the diffuser ring are bolted the cast elbows conducting the air to the combustion chambers. In the bend of the elbows are three cascade vanes formed of lengths cut from an extruded seetion and cast in position. A pair of trunnions providing the main supports for the complete unit are also mounted on the diffuser ring. The front bearing housing forms the outer member of the front air intake and also serves to support the wheelcase containing the auxiliary drives and the oil pump. Bolted up to the turbine shroud ring and nozzle box is the exhaust with its inner cone supported by four transverse bolts enclosed by strearnlined fairings. The base of the inner cone masks the rear face of the turbine disc. While the exhaust cone is of fixed length, approximately 33in., the jet pipe extending from the exhaust cone to the propulsion nozzle can be varied to meet installation requirements providing a suitable length/ diameter ratio is maintained. These parts are double walled and packed with heat insulating material. Standard length of jet pipe is about 44in. and the weight is 9.5 Ib. per foot. Combustion chambers are similar in general design to those of the Derwent V but of larger capacity. Nene units at present in production are rated at 5,000 Ib. thrust. This is neither the maximum at present available nor the ultimate possibility. A thrust of 5,500 Ib. has already been obtained on the test bed. Average figure during development was 5,150 Ib. which represents the following component efficiencies :- Compressor 76 per cent., Combustion 98 per cent., Expansion (turbine and tail cone) 93 per cent. A compressor having a double-sided impeller was chosen because output of a jet unit is largely determined by the amount of air consumed and this is conditioned by diameter of the compressor entry. Obviously two intakes will admit more air than one. Conversely, for a given quantity of air, the overall diameter of the impeller and consequently the complete unit can be relatively smaller than on a single-sided design, which is advantageous particularly with wing installations. There are, of course, other reasons influencing the choice. The increased air flow for any given diameter necessitates relatively long turbine blades and a smaller diameter turbine disc and permits an advantageous stressing of these parts. A wing nacelle or a fuselage enclosure forms a plenum chamber from which air is drawn to the compressor intakes. Velocity is lowered and any object sucked in may well fall to the bottom of the enclosure instead of passing into the compressor. The first aircraft to be powered by the Nene was a Lockheed XP80 Shooting Star and recently tests have been conducted on a De Havilland Vampire. In both instances an improvement in performance was obtained. With the American aircraft speeds of the order of 580 m.p.h. and an excellent rate of climb to 42,000ft. were achieved.

    G. Geoffrey Smith M.B.E.,
    Gas Turbines and Jet Propulsion for Aircraft, 4th ed. 2nd imp., Oct 1946.
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