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Aviation History
1957
1957 - 1078.PDF
168 FLIGHT, 2 August 1957 •- „•. .. • I (Sketch, left) Operation of the variable-incidence inlet guide vanes. The actuator ring—supported on a row of bearings—carries a series of spherical bearings locating the vane swivel arms. (Sketch, right) This typical compressor detail shows the shrouded stator blade roots and the hollow swivel-pin attachment of the compressor blades. Inter-stage sealing is effected by a controlled clearance between individually machined brass labryinth seals and the compressor rotor drums. The formed sheet first-stage nozzle blades house the tie rods (extreme right) supporting the inner shroud ring and rear main bearing. Second-stage nozzle blades are pin-jointed to their shroud rings. GY RON . . . from solid Nimonic 80A forgings and have a single-tang anddouble-tang fitting at their inner and outer ends respectively. The blades are pin-mounted to the inner and outer shroud rings viathese tangs and are assembled with an offset angle from a truly radial position; in this way expansion of the blade length is accom-modated by slight rotation of the inner ring. Such an arrange- ment has the additional advantage that a certain proportion ofthe gas bending-loads on the nozzles is taken in tension rather than in pure cantilever bending. A sheet-metal labyrinth sealis mounted off the inner shroud ring. This acts in conjunction with the inter-stage drum to prevent axial leakage of the gasesbetween the two stages. During development testing of the engine, fatigue failures of theroot profile form and of the serrations occurred in both turbine rows. These were determined as being caused by sixth-orderexcitation in the first stage and third-order excitation in the second stage. It was for this reason that the number of fuel injectors inthe combustion system was altered from eighteen to seventeen. The serration failures were caused by stress concentration set upby inequality of stress-flow from the blade platform into the root. After a considerable amount of design and test work, it wasconcluded that for engines rated at 15,000 lb—and in particular those intended for flight testing—the best compromise to combatfatigue conditions was the use of shrouded Nimonic 90 first- and second-stage blades. Conversely, for the higher-rated, higher-temperature engines, it was necessary to replace the Nimonic 90 first-stage blades with buttressed, unshrouded Nimonic 100 blades.These latter have a much thicker profile near the blade platform and are better suited to resist fatigue conditions at the highertemperatures. A straightforward design of exhaust cone assembly fabricatedlargely from Nimonic 75 sheet is fitted. The construction of this is apparent in the cutaway drawing. Air System. A tapping of air for cooling purposes is bled froma ring of holes in the inner casing of the combustion chamber. A portion of this air is directed over the front face of the first-stageturbine disc and escapes into the gas stream at the periphery of the disc. A further portion of the air passes through a series ofholes drilled in the first-stage disc and flows under the inter-stage drum. Slots cut in the rear flange of this drum permit the air toescape into the gas stream at the periphery of the second-stage disc. The remainder of the cooling air passes up the first stagenozzle blades and thence between the pairs of turbine and nozzle outer shroud rings. Fuel System. The complete fuel system on the Gyron hasbeen supplied by Joseph Lucas (Gas Turbine Equipment), Ltd., who also largely developed the annular combustion system and theatomizing fuel-injection system—the first to be fitted to a British engine with a large annular chamber. The fuel system comprisestwo G.D. size multi-plunger, variable-delivery pumps, a propor- tional-flow control unit and a flow distributor. Provision is alsomade for electric speed and temperature over-ride trims. The proportional control unit consists of a throttle valve/shut-off cockand Unking valve forming the flow transformer, an acceleration control and constant-idling-flow valve. A separate assembly con-tains the altitude-sensing unit, which, with die flow transformer, forms the scheduling control. A further unit, the electro-pressure control, contains thesecondary-flow control orifice and also receives amplified signals from an engine-driven alternator and the jet pipe thermocouples.The signals vary the secondary flow and so provide speed and temperature trim to the scheduling system. Lubrication System. A dry-sump system supplies oil (spec.D.Eng.R.D. 2487) under pressure to the front bearing, centre- housing gears and bearings, and rear bearing. Oil from the frontbearing and centre housing drains back into the sump via the lower intake strut; oil is retained in the rear bearing by piston-ring sealsfore and aft of the rollers. A pump passes scavenge oil to the top of the integral oil tank around the intake casing. A fuel-cooledoil cooler (not shown on the sectional drawing) is proposed for cooling the oil under high forward speed conditions. Starting System. The engine is started by either a 45 h.p.110 volt D.C. Rotax electric starter motor or a de Havilland hydrogen peroxide turbo-starter. The starting times that can beachieved by these two systems are known to be in the order of 30 sec and 10 sec respectively. The starter is linked to the mainrotative assembly by reduction gearing (each system has a particu- lar reduction ratio) and an extension shaft from the front of thecompressor rotor. Constituting a highly important aspect of the overall pro-gramme on the Gyron has been the research and development work performed at the Halford Laboratory. This establishmentis extensively equipped with the necessary facilities for rig test- ing of the major components of the engine and the related intakeand exhaust systems. Because of the large size of the Gyron and the very high powersand air mass flows involved, certain of this work has been more conveniently performed on components of the scaled-down GyronJunior. Examples include the compressor and compressor stage testing mentioned previously, together with atmospheric tests ofthe combustion system and cold-flow tests of the Junior turbine. Information obtained from static and rotating rig tests of GyronJunior cooled turbine blades has also been used in the design of blades for the larger engine. There is sufficient similarity between the two designs of engineto enable further comparisons to be made. By using the com- pressor test rig in the Halford laboratory, supersonic ram airconditions can be applied to a Gyron Junior on an adjacent test- bed and should provide accurate information applicable to theGyron engine family. Various fundamental types of supersonic intake are being investigated by a programme of model research.Facilities are now available for testing small-scale intakes up to Mach 2, but plans are in hand by the de Havilland engine com-pany for a 24in-square-section tunnel capable of operating at Mach 3—the flight conditions intended for the Gyron. Development of this large, ultra-powerful engine is witnessto the forward thinking of its makers in their determination to ensure that a developed supersonic engine is available in advanceof the requirements of the aircraft manufacturer. In particular, the mixed rocket-plus-turbojet powerplant formula has beenclosely studied and the Gyron has been considered for installa- tion in the projected designs of a long-range supersonic bomber,supersonic fighter and cruise missile. With the re-shaping of the national defence policy announcedearlier this year in the Government's White Paper on Defence, it was declared that with the evolution of guided missiles for bothoffensive and defensive roles the Gyron (after over four years of development work and 2,000 hours of bench and flight testing)would no longer receive Government financial support. This might have sounded the death knell of a less imaginative or a lesssuccessful project, but the practical interest shown in the engine by the industry—in particular by the Hawker Aircraft Company(who, it is announced, are to use the Gyron in their P.1121 prototype supersonic aircraft)—decided the de Havilland EngineCompany to proceed with the development on a private venture basis as before. Adequate backing, they say, is being provided forflight engines for the P.1121 and other forthcoming aircraft. •'••-J--.'!&£.•-•-: "
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