Concorde's Olympus Engine

The Rolls-Royce Olympus (originally the Bristol B.E.10 Olympus) was the world's first two-spool axial-flow turbojet aircraft engine, originally developed and produced by Bristol Aero Engines. First running in 1950, its initial use was as the powerplant of the Avro Vulcan V bomber. The design was further developed for supersonic performance as part of the BAC TSR-2 programme. In its final version designed for aircraft, of course it saw production as the Rolls-Royce / Snecma Olympus 593, the powerplant for Concorde

At the end of World War II, the Bristol Engine Company’s major effort was the development of the Hercules and Centaurus radial piston engines. By the end of 1946, the company had only 10 hours of turbojet experience with a small experimental engine called the Phoebus which was the gas generator or core of the Proteus turboprop then in development. Bristol Aero Engines (formerly Bristol Engine Company) merged with Armstrong Siddeley Motors in 1959 to form Bristol Siddeley Engines Limited (BSEL) which in turn was taken over by Rolls-Royce in 1966.

The thrust required of this new engine, then designated the B.E.10 (later renamed the Olympus), would initially be 9,000 lbf (40kN) with growth potential to  12,000 lbf (53kN). The pressure ratio would be an unheard of 9 to 1. To achieve this, the initial design used a low pressure (LP) axial compressor and a high pressure (HP) centrifugal compressor, each being driven by its own single-stage turbine. This two-spool design made the compression more manageable, enabled faster engine acceleration ("spool up"), and reduced surge. The design was progressively modified and the centrifugal HP compressor was replaced by an axial HP compressor. This reduced the diameter of  the new engine to the design specification of 40in (100cm).

The first engine, its development designation Bristol  Olympus 1, had six LP compressor stages and eight HP stages, each driven by a single-stage turbine. The combustion system was novel in  that ten connected flame tubes were housed within a cannular system: a hybrid of separate flame cans and a true annular system. Separate combustion cans would have exceeded the diameter  beyond the design limit and a true annular system was considered too  advanced.

The Rolls-Royce / Snecma Olympus 593 as used in Concorde was an afterburning (reheated) turbojet which was initially a joint project between Bristol Siddeley Engines Limited (BSEL) and Snecma. Until regular commercial flights by Concorde ceased, the Olympus turbojet was unique in aviation as the only afterburning turbojet  powering a commercial aircraft. The overall thermal efficiency of the engine in cruising flight was  about 43%, which was the highest figure recorded for any normal  thermodynamic machine. A quieter, higher thrust version, the Mk 622, was proposed where reheat was  not required and the lower jet velocity reduced the noise from the  exhaust. The improved efficiency would have allowed greater range and opened up  new routes, particularly across the Pacific as well as transcontinental  routes across America. However, the poor sales of Concorde meant that  this plan for a Concorde 'B' model fitted with this engine was not pursued.

Development of the engine and engine accessories was the  responsibility of Bristol Siddeley, while the Snecma company was responsible for the variable intake, the exhaust nozzle / thrust reverser / noise attenuation and the afterburner. Britain was to have a larger share in production of the Olympus 593 as France had a larger share in fuselage production. The Olympus 593 was a 2-shaft turbojet with reheat. The LP and HP  compressors both had 7 stages and were each driven by a single stage  turbine. The compressor drums and blades were made from titanium except  for the last 4 HP stages which were nickel alloy. Nickel alloys were normally only required in the hotter turbine areas  but the high temperatures that occur in the last stages of the  compressor at supersonic flight speeds dictated its use in the  compressor also. The HP turbine rotor blades were cooled. A partial reheat (20% thrust boost) was installed to give the required take-off thrust. It was also used  for transonic acceleration from just below Mach 1 up to Mach 1.7; the  engine supercruised above that speed and at cruise the thrust through the engine mounts contributed 8% of the thrust from the complete propulsion system.

The final version, as fitted to Concorde, was the Olympus 593-610-14-28 which produced 32,000 lbs force (142kN) of thrust without reheat and 38,050 lbs force (169kN) of thrust with reheat.