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Aviation History
1956
1956 - 0019.PDF
6 January 1956 19 THIRTEEN-STASE COMPRESSOR 1 Cartridge-starter breeches. 2 Rotax cartridge starter. 3 Starter reduction gearbox. 4 Starter engagement clutch. 5 Internal wheel case. 4 Accessory-drive coupling. 7 Drive to h-p pump and tachometer. 8 Intake casting. • Strut leading edge. 10 Inlet extension piece. 11 Integral front mounting. 12 Tank breather pipe. 13 Overboard breather pipe. 14 Forward lifting eye. 15 Oil scavenge line. 14 Main oil pump. 17 Micropump (centre bearing). 18 Overspeed governor head. If Fuel by-pass valve. 20 Dowty fuel pump. 21 Installation trunnic22 "Banana" wheel case. 23 Tachometer. 24 Fuel to l-p inlet. 25 L-p filter and flow-control. 24 Distributor (one of six). 27 Fuel and air drain and oil breather. 28 Outlet for bearing-cooling air. 2f Fuel feed (one of thirty-six). 38 Primer jet (one of six). 31 High-energy igniter (one of a pair). 32 Combustion chamber drain.33 Front main shaft. 34 Oil pump drive. 35 Fixed guide vanes. M Locating shoulders on rotor. 37 Holes for eighth-stage air. 38 Tapping for fifth-stage air (one of six) 39 Seventh-stage air for gun heating or cooling. 48 Fuel manifold. 41 Straightener vanes. 42 Centre-section. 43 Slinging unit. 44 Bearing oil down fifth-stage air pipes. 45 Outer flame tube. 44 Primary air tube. 47 Feed-pipe built up to fill hole. 48 Secondary air tube. 4* Primary air tube held by dogged- periphery nut with crimping washer. 50 Primer ring. 51 Spherical bearing. 52 Stator conical support unit. 53 Drillings to admit turbine cooling air. 54 Rear (fixed) cone. 55 Front (split) cone. 54 Front cone retaining nut« 57 Turbine lock nut. 58 Sprung plunger. 59 Sliding gas seal. 40 Torque-transmitting dogs. 41 Cooling air passage around first stacors. 42 p4 governor line. 43 Turbine locking dowel. 44 Metering plug. 45 Exhaust cone shroud. 44 Alfoi lagging. 47 Disc with pressure-balance holes. 48 Welded internal stiffeners. 49 Spring latches. 78 Pipe attachment access door. made to steel blading, also riveted and of identical profile,although further failures across the rivet holes persisted. Different types of root were investigated in conjunction withblades of increased thickness /chord ratio. The final choice fell upon a type of blade with a serrated (fir tree) root held in arelatively thick, steel disc, and this was the arrangement adopted for the first three rotor stages. Further running resulted infatigue failures through the serrations but this was completely cured by increasing the root cross-section. In common with several other axial engines of low hub/tipratio the first few stages of the Sapphire compressor operate in a semi-stalled condition at low r.p.m. and this resulted in thefatigue failures mentioned above. One of the attempted solu- tions was the lacing together of the rotor blades in the first stageby a pre-formed wire threaded through small holes at approxi- mately mid-chord near the tips. The presence of this lacing wiresufficed to damp out any resonance. Later development, how- ever, made such measures unnecessary. In October 1950, a licence agreement was concluded betweenthe Wright Aeronautical Division of the Curtiss-Wright Corpora- tion and Armstrong Siddeley, under the terms of which theAmerican company were enabled to produce the Sapphire in an "Americanized" form. At this time the principal British enginewas the Sa.3, and several examples were sent to the New Jersey firm during 1950 and 1951. The history of the Sapphire in theU.S.A. is fully covered later in the course of this narrative. The only other difficulties experienced with the Sa.3 wererelatively trivial failures of the type met in all engine develop- ment programmes. For example, the cooling effect of eighth-stage bleed air between the two turbine discs was insufficient, and failures of the locking band occurred. A change to 13th-stage airprovided a completely adequate cooling flow. Another source of trouble was the inlet guide vane assembly, in which the bladeswere brazed at their inner ends to 16 shroud segments, each segment holding three blades. Fatigue failures of this section werestopped by making the shrouds in complete 180-deg segments. Similar experience caused a change in design of the brass com-pressor stator-blade shrouds; originally in eight segments per stage, they were redesigned in 90-deg quadrants. Throughout 1951 detail improvements of this nature continued,and in January, 1952, the Sa.3 passed a second M.o.S. type test at 7,590 lb thrust and with slightly better fuel consumption (0.909). From the outset one of the chief assets of the Sapphire wasits ability to swallow a large airflow in relation to the overall diameter of the engine. Nevertheless, the extended bench testingcarried out during 1951 with the Sa.3 showed that an even higher mass flow could be accepted, with slight modification to the gaspath—such as a reduction in the swirl angle of the inlet guide vanes—merely by running the engine at higher r.p.m. Develop-ment became concentrated upon this up-rated engine, designated ASSa.6, and culminated in a remarkably fine type test in March,1952, at no less than 8,300 lb thrust and with an s.f.c. of 0.9. High-altitude testing began in August 1951, when twoSapphires, of the Sa.3/Sa.6 type (the latter rating had not then been achieved) were flown in WD 933, the fifth production Eng-lish Electric Canberra. This was one of the first of many Canberra engine test beds and it has since carried much of the load ofSapphire testing on its broad wings. As is now self-evident, the Korean conflict caused a rapidincrease in procurement for all types of military equipment in Britain and America, and the Sapphire was ordered in quantityearly in 1950. The first production order referred to engines of the Sa.3 rating but the Sa.6 development rapidly overtook thetooling-up process, with the result that the contract was changed
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