Rolls-Royce/Snecma Olympus 593

The initial design of the engine was a civil version of the Olympus 22R, redesignated as the 591. with considerable improvement in performance. Development of the engine and engine accessories was the responsibility of Bristol Siddeley; BAC was responsible for the variable intake and overall engine installation, with Snecma taking on the exhaust nozzle incorporating the thrust reverser, noise attenuation, and reheat. Britain was to have a larger share in production of the Olympus 593 based powerplant as France had a larger share in fuselage production. Ground test running of the engines was co-ordinated between Bristol Siddeley, Patchway; the National Gas Turbine Establishment (NGTE), Pystock, UK; and the Centre d'Essais des Propulseurs (CEPr) at Saclay, France. The Olympus 593B was first run in November 1965. The B (for "Big") was a redesign of the 593D which was planned for an earlier smaller Concorde design. Test results from the 593D were used for the design of the B. The B was dropped later from the designation. Snecma used an Olympus 301 in testing scaled models of the nozzle system. In June 1966, a complete Olympus 593 engine and variable geometry exhaust assembly was first run at Melun-Villaroche. At Bristol, flight tests began using a RAF Avro Vulcan bomber with the engine and its nacelle attached below the bomb-bay. Due to the Vulcan's aerodynamic limitations, the tests were limited to a speed of Mach 0.98 (1,200 km/h). During these tests, the 593 achieved 35,190 lbf (157 kN) thrust, which exceeded the specification for the engine. In early 1966, the Olympus 593 produced of thrust with reheat. In April 1967, the Olympus 593 ran for the first time in a high altitude chamber, at Saclay. In January 1968, the Vulcan flying test bed logged 100 flight hours, and the variable geometry exhaust assembly for the Olympus 593 engine was cleared at Melun-Villaroche for flight in the Concorde prototypes. Concorde prototype 001 made its maiden flight from Toulouse on 2 March 1969. It was captained by André Turcat, chief test pilot of Sud Aviation. Using reheat it lifted off at 205 knots (380 km/h) after a ground run of 4,700 feet (1.4 km). A quieter, higher thrust version, the Mk 622, was proposed. 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' was not pursued. ==Propulsion system design==

Engine The Olympus 593 was a two-shaft turbojet with reheat. The low pressure (LP) and high pressure (HP) compressors both had seven stages and were each driven by a single-stage turbine. Due to the high inlet air temperatures at Mach 2 cruise - in excess of 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 stator and rotor blades and LP turbine rotor blades were cooled. Partial reheat providing a 20% thrust increase File:Olympus 593 AB.jpg|Cutaway Olympus 593 showing LP compressor, HP compressor, combustor, turbines and reheat fuel ring and flame holder. File:CONCORDE AIRCRAFT ENGINE MATERIALS - NARA - 17446865.jpg|Olympus 593 drawing showing two shafts and reheat. Materials required for high intake temperatures at cruise are: Green- steel alloys, Red - Waspaloy, Yellow - titanium alloys, Blue - nickel-based alloys. Partial reheat only required a single fuel ring and flame stabiliser as shown. File:Olympus 593 FAAM Turbine.JPG|Olympus 593 view of partial reheat (20% thrust boost) showing lower fuelling requirement/temperature, needing only a single ring, compared to a military full reheat. File:F104Engine.JPG|For comparison with the 593, a military full reheat requirement (50% thrust boost/3,000 degF) needing three fuelling rings (hidden by red flameholders) for flame across the whole duct (GE J79 turbojet engine). Intake The Concorde's variable-geometry intake, designed by BAC, like any jet-engine intake, has to deliver the air to the engine at as high a pressure as possible (pressure recovery) and with a variation in pressure distribution (distortion) that can be tolerated by the compressor. Poor pressure recovery is an unacceptable loss for the intake compression process, and excessive distortion causes engine surging (from loss of surge margin). If the engine is a turbojet with reheat, the intake also has to supply cooling air for the hot reheat duct and engine nozzle. Meeting all the above requirements over the full range of the certified operating envelope was required for Concorde to become a viable commercial aircraft. They were met with a variable geometry intake and an intake-control system that compromised neither the operation of the engine nor the control of the aircraft. Supersonic pressure recovery is addressed by the number of shock waves that are generated by the intake: the greater the number, the higher the pressure recovery. Supersonic flow is compressed or slowed by changes in direction. The Concorde intake front ramps changed the flow direction, causing oblique external shocks and isentropic compression in the supersonic flow. The TSR-2 had used a half-cone translating centre-body to change the direction. Since the ramp bleed slot was in the subsonic diffuser, and downstream of the shock system, changes in flow demanded by the engine would be accommodated with corresponding changes in the bleed slot flow without significantly affecting the external shock pattern. Engine flow reductions caused by throttling or shutting down were dealt with by dump-door opening. The primary exhaust nozzle and jet pipe were designed for a life of 30,000 hours; the Thrust Reverser Aft (TRA) structure for a life of 40,000 hours. File:Concorde - Imperial War Museum - Duxford - Cambridgeshire - England (27990675820).jpg|The pre-production Concorde G-AXDN, Duxford, close-up of nozzles and reverse-thrust, exit-air cascade vanes. On production aircraft, a revised design of "eyelid" variable nozzle/thrust reverser was used. File:Rolls-Royce-Snecma Olympus - Musée Safran.jpg|Engine and production nozzle arrangement under one Concorde wing. The monobloc nozzle assembly is installed behind two Olympus 593s. The nozzles are shown in different operating positions to illustrate take-off (right) and landing reverse thrust (left). They are partially closed for take-off to reduce sideline noise. They are fully closed to deflect the exhaust partially forward to produce reverse thrust. File:ConcordeIntakeNozzle.svg|Airframe secondary nozzle positions when A) taking off B) Mach 2 cruise C) thrust reversing. ==Variants==

• 593 - Original version designed for Concorde • Thrust : 20,000 lbf (89 kN) dry / 30,610 lbf (136 kN) reheat • 593-22R - Powerplant fitted to prototypes. Higher performance than original engine due to changes in aircraft specification. • Thrust : 34,650 lbf (154 kN) dry / 37,180 lbf (165 kN) reheat • 593-610-14-28 - Final version fitted to production Concorde • Thrust : 32,000 lbf (142 kN) dry / 38,050 lbf (169 kN) reheat ==Engines on display==

Preserved examples of the Rolls-Royce/Snecma Olympus 593 are on display at the following museums: • Aerospace Bristol, Bristol, UK • Newark Air Museum, Newark, UK • Technik Museum Sinsheim, Sinsheim, Germany In addition to these museums, other sites that display examples of the Olympus 593 include: • Talbot Laboratory at the University of Illinois at Urbana-Champaign, Illinois, US • Weber Space Science and Technology Building at the Georgia Institute of Technology, Georgia, US • Whittle building at the Cranfield University, Cranfield, UK File:Rolls-Royce-Snecma Olympus - Musée Safran.jpg|An engine pair with thrust reversers at the Musée aéronautique et spatial Safran File:RR-Snecma Olympus 593-602B.jpg|An engine on display at Cranfield University ==Specifications (Olympus 593 Mk 610)==

section of an Olympus 593 on display at the Fleet Air Arm Museum • Overall pressure ratio: about 82:1 at Mach 2.0 (includes pressure rise in intake, 7.3:1, and in engine, 11.3:1 which is less than when stationary due to ram temperature rise) }} ==See also==