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
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.
version 2015-01-19.
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