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Aviation History
1957
1957 - 1073.PDF
"--:J> 2 August 1957 GYRON de HavillancTs Supersonic Turbojet, with 25,000 Ib. Reheat Rating A LTHOUGH four years have passed since it was announced/\ that the de HavUland Engine Company had in existence a L JL. truly supersonic turbojet, until very recently only conjec-tural analysis of the Gyron's design features could be made. The following full description and cut-away drawing emphasize thesimplicity of this very powerful engine and record some of the solutions that have been adopted—as a result of over four yearsof rig and flight development—to the problems of developing a supersonic gas turbine. The decision from which die Gyron engine stemmed wasarrived at seven years ago, when the company made a detailed consideration of the powerplant to succeed the Goblin and Ghostcentrifugal-compressor engines. At this time they initiated a painstaking and comprehensive programme of exploration intothe performance characteristics of supersonic turbojets, the aero- dynamic and thermodynamic considerations governing the designof such an engine, and the related aspects of the powerplant as a combustion system. Not the least of the problems to be facedat this stage, and quite apart from those of design and manufacture, were associated with component development and with groundand flight testing. A year later, the further decision to proceed with the design ofa high-thrust turbojet for use at Mach 2-3 on a company-financed basis was taken under the direction of the late Major FrankHalford, who was then chairman and technical director of the Engine Company. As no corresponding decision had then beenmade within either the Government or the industry to proceed with an associated design of aircraft, this was a bold and far-sighted step to take, deeply involving the company's financial resources. It meant that de Havilland engineers were designingfor powerplant requirements of a period of, probably, five to fifteen years ahead, and maximum use was made of the compre-hensive theoretical investigations already undertaken. It had been determined that, because of the high flight-speedsconcerned, the intake system assumed a greatly enhanced import- ance. The air approaching the intake exhibited a high level ofkinetic energy suitable for conversion to pressure energy by the intake; and the magnitude of the ram pressure available at theintake face of the engine was such that if the optimum pressure- ratio in the engine cycle was to be utilized it was only necessaryfor the compressor to achieve a moderate pressure ratio. In these circumstances (i.e., where the major proportion of the overallpressure-rise in the engine cycle could be obtained in the aircraft intake) it was obvious that this component should be designed toperform its duty of diffusing the incoming air with the maximum efficiency. Likewise it had been determined that if very high flight speedswere to be attained a much hirfier efflux velocity—and hence higher thrust output—was essential. This could be achieved withoutunduly diminishing the propulsive efficiency of the aircraft. In addition, the improvement in engine specific thrust (i.e., lbthrust/lb air mass flow/sec) and fuel consumption could be achieved by increasing the turbine inlet temperature. . The radically higher overall pressure ratio of a supersonicengine implies a similarly increased expansion ratio across the exhaust system. As this latter ratio is virtually constant, theexpansion ratio across the propelling nozzle rises with increase in flight speeds; and if this available expansion ratio is to be used toits maximum efficiency a convergent-divergent nozzle capable of achieving supersonic expansion of the exhaust gases is necessary. To assess the precise trend of these various effects, curves ofspecific thrust and specific fuel consumption, plotted in terms of compressor pressure ratio and turbine entry temperature, wereproduced for a range of flight conditions. They showed that in the consideration of the highest turbine temperatures likely tobe practicable in the near future, a particular compressor pressure ratio (a reasonable assumption is 5 or 6 : 1) represents the optimumfor supersonic propulsion in the stratosphere. Design work on the new engine, designated the H.4 as thefourth of die de Havilland/Halford turbine engines, started late in 1951. Subsequently named the Gyron, it was conceived as asimple, lightweight single-shaft unit having a design thrust rating of 15,000 lb)—a figure very much in excess of that of any otherengine known to be under consideration at the time. The basic layout involved a seven-stage axial compressor and a two-stageturbine united by a large-diameter shaft, the complete rotative assembly being carried on two main bearings only. The combus-tion system was of fully annular design and employed upstream fuel injection. Simultaneously with the manufacture of parts for the prototypeengine^ additional main components were produced for individual testing on rigs. In particular, a rig compressor was built fortesting at C. A. Parsons, Ltd., some months before the initial run of the engine, and an open-type test-bed formerly used for benchtesting of the Ghost was specially converted to provide for the considerably greater dimensions and thrust of the Gyron. It was on this bed that the Gyron DGy.l first ran on January 5,1953. Apart from minor adjustments which it was found neces- sary to make to the setting of the inlet guide vanes, design per-formance was quickly achieved; very high thrust figures were in fact attained by the Gyron within a few weeks of the initiation oftesting. And in April of the same year, after the capabilities of the engine had been physically established, it was awarded a develop-ment contract and official financial backing by the Ministry of Surroly. During the first year of its development a Gyron with modifiedturbine blading, combustion and fuel systems completed a 25-hour soecial-category test at 13,000 lb in May, and by October morethan 150 engine hours had been accumulated, with several hours performed at design thrust. In July 1954, a 25-hour run wasundertaken at 15,300 lb and this was followed by further runs in which the engine maximum temperature and r.p.m. were pro-gressively increased. In this manner a thrust of 16,500 lb was initially achieved, to be followed by over 100 hours' running at17.800 lb thrust. Then in September, in view of the continued satisfactory state of the engine, several runs were undertaken at
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