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Aviation History
1957
1957 - 1077.PDF
GYRON . . . segments are shrunk into a circumferential steel carrier ring whichis attached to the inside of the forward entry duct and spring- loaded to permit a limited radial movement. With the enginerunning, the rotating labyrinth moves the carbon seal to a con- centric position, and only a slight rubbing of the carbon occurs.As engine r.p.m. are increased the higher compressor air delivery pressure holds the seal firmly against any further movement. Someof the original seals are still in use on Gyron development engines. To supplement information on compressor performance gainedduring engine running, rig testing of individual compressors was undertaken. This work was performed on a 2,500 h.p. closed-circuit rig at C. A. Parsons, Ltd., and initially a six-stage unit was built and tested. During 1954 and 1955 further testing of anintensive nature was performed on a seven-stage compressor at the National Gas Turbine Establishment and a useful measure ofoverspeeding was also undertaken. The performance results obtained confirmed those predicted during design. At the Company's new gas dynamics laboratory at Hatfield,named the Halford Laboratory, extensive rig testing of a thin- bladed lightweight design of compressor was completed earlyin 1956. Performance investigations have similarly been made of special single-stage compressors and of compressors havingvariable incidence first-stage stator blades. These tests, and the handling of the engine on the test-bed and in flight, have shownthe Gyron to be capable of an exceptionally high rate of stall-free acceleration—three seconds from flight idling to full thrust. In addition to the thin-bladed compressor, a lightweighttitanium-bladed compressor has also been tested. The rotor of this unit incorporated four discs also machined in titanium alloy.The engine concerned was tested to a high rating and with reheat developed 23,000 lb thrust. The forward entry duct, interposed between the compressorand the combustion system, forms the backbone of the engine. In view of the ambient temperatures to which it is subjected, itis produced as a fabricated structure in Fortiweld steel and com- prises two double-walled drums forming the diffuser outlet fromthe compressor. Seventeen streamlined struts have the dual role of carrying the loads from the inner drum to the outer drum and ofsupporting the flame-tube assembly at its forward end. Bolted to the rear of the inner drum is a further pair of drumsfabricated in Fortiweld. The outer is a non-stress-carrying item which defines the inner wall of the combustion system, and theinner drum is slightly conical and extends as far as the front of the turbine entry duct. At this point a further conical drum,also in Fortiweld, extends rearwards to carry the rear main bear- ing housing. A large flange at the forward end of this secondconical drum provides attachment spigots for the inner portion of the turbine entry duct and for a fabricated conical baffle unitwhich is provided for directing the flow of cooling air over the turbine disc. The combustion chamber of the Gyron was the greatestunknown in the design of the engine. In its original form it comprised an annular snout, a swirler-plate assembly and an innerand outer flame-tube, each in Nimonic 75. The outer air casing in Fortiweld steel was also a structural load-carrying member;to the inner air casing reference has already been made. Eighteen Lucas Duplex 3 burners, providing upstream fuel injection, wereused. Reversing air chutes were provided part-way down the flame-tube walls to inject air forwards into the primary combustion zone between adjacent fuel injectors. It was the recirculationof airflow created by these and the swirler air streams that effec- tively stabilized and anchored the position of the flame. A seriesof corrugated strips were provided at stages down the flame tube wall to pick off layers of dilution air to skin-cool the entire innerface of the flame-tubes. The exceptionally high mass flow of the engine prohibited anypossibility of providing an adequate air supply to test the chamber at full design conditions, but testing of the chamber on an FLIGHT, 2 August 1957 167 (Left) The first-stage nozzleassembly utilizes hollow blades, for both structural and cooling reasons. (Right) Familiar fir-tree serrationsare used to attach the turbine blades to their discs. The second-Stage assembly shown here is attached to the first by an inter-stage drum and is spigotted to the mainshatt. atmospheric pressure rig was performed at the N.G.T.E. at thesame time as the Gyron was first run. Subsequently, a large num- ber of similar tests have been undertaken at the Halford Labora-tory. The initial results of this work and early development running on the engine indicated that an unsatisfactory turbineentry temperature pattern prevailed. Several re-designed forms of chamber were tested, including an eleven-burner configurationhaving convoluted inner and outer flame-tubes with the object of obtaining a more uniform supply of air to die fuel sprays. As a result of this work, modifications were made to thedilution air ports. When tested on an eighteen-burner chamber in an engine these indicated that an appreciable improvement inturbine entry temperature conditions had been obtained. It was largely this change which enabled die attainment of 20,000 lbthrust by the Gyron DGy.2. The final choice of a seventeen- burner arrangement was made as a means of avoiding certainresonant conditions in the turbine blading. Subsequent extended running at high engine thrust showedthe need for a further improvement in the radial temperature gradient at entry to the turbine, to prevent stretching of thefirst-stage turbine blades. It was determined that changes were required in the primary-zone burning conditions. These wereachieved by the introduction of ducted swirlers rather than the more conventional single annular snout. This latter componentwas modified to split the incoming airflow from the compressor into two streams. From the inner one of these, primary air forcombustion is individually ducted to the seventeen swirler plates, thus providing a more uniform and direct supply of air for thispurpose. Joining the compressor rotor to the turbine is the forged S.97steel mainshaft, machined to an exceptionally thin wall-thickness over the major portion of its length. The trumpet profile of thiscomponent was developed to provide a shape capable of resisting the centrifugal loads in the shaft at high rotational speeds. Theshaft is supported at its rear by the main single-row roller- bearing assembly and stator casings. To ensure that the two turbine discs are held concentric withone another and with the mainshaft, it is arranged that the first- stage disc is located on the shaft by means of a Hirth coupling,while the second-stage disc is attached to the first by means of an inter-stage drum fitted with dowels. The forward portion of theHirth coupling is carried on splines over the rear of the mainshaft and the rear half of the coupling is machined directly on the frontface of the first disc. Radial growth of the disc is permitted by the radial disposition of the teeth of the coupling; and, to ensurethat adequate flank contact of the teeth is maintained under all flight conditions, a large ring-nut abuts against the rear face of thefirst-stage disc. The first-stage disc is a forging in Firth-Vickers Rex 448,and the second-stage disc, which is also spigotted at its centre to the rear of the main shaft, is a forging in William Jessops H.40.Both of these materials are heat-resisting ferritic steels. The turbine blades are fully machined from Nimonic 100 forgings forthe first stage and Nimonic 90 for the second stage, and are retained in the discs by five-serration fir-tree root fittings.Nimonic 75 rivet pins at the apex of the fir-tree fitting lock the blades against endwise movement. The first-stage nozzle blades are formed to shape from Nimonic80A sheet, seam-welded along the trailing edge. The tie-rods previously mentioned pass through these to die outer shroud ring,which also carries the engine rear-support linkage. To accom- modate radial expansion and blade torque reaction, the nozzlesare carried in profiled slots cut in the inner and outer shroud rings. The inner shroud also forms one side of an annular boxwhich is supplied with a bleed of compressor delivery air passing inside die blades to cool them. Both the turbine and nozzle bladesare surrounded at their tips by inner and outer shroud rings cooled by the rearward flow of air passing between the rings before itexhausts into the gas stream at the trailing tip of the second-stage blades. The second-stage nozzle blades are of necessity cantilever-mounted from their outer shroud ring. They are fully machined
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