Rolls-Royce/Snecma Olympus 593

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Olympus 593
Olympus593.JPG
Preserved Olympus 593 engine at the Imperial War Museum Duxford
Type Turbojet
National origin United Kingdom/France
Manufacturer Rolls-Royce Limited/Snecma
First run June 1966
Major applications Concorde
Number built 67
Developed from Rolls-Royce Olympus

The Rolls-Royce/Snecma Olympus 593 was an afterburning (reheated) turbojet which powered the supersonic airliner Concorde. It was initially a joint project between Bristol Siddeley Engines Limited (BSEL) and Snecma. It was based on the Bristol Siddeley Olympus22R engine.[1] Rolls-Royce Limited acquired BSEL in 1966 during development of the engine making BSEL the Bristol Engine Division of Rolls-Royce.[2]

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.[3]

Development[edit]

The initial design of the engine was a civil version of the Olympus 22R, redesignated as the 591.[1] The 22R had been designed for sustained (45 minutes) flight at Mach 2.2[3] as the engine for the BAC TSR-2. The 591 was redesigned, being known as the 593, with specification finalised on 1st January, 1964.[1] Bristol Siddeley of the UK and Snecma Moteurs of France were to share the project. SNECMA and Bristol Siddeley were also involved in an unrelated joint project, the M45H turbofan.

The early development stages validated the basic design concept, but many studies were required to meet the specifications which included fuel consumption (SFC), engine pressure ratio, weight/size and turbine entry temperature.

Initial studies looked at turbojets and turbofans, but the lower frontal cross-sectional area of turbojets in the end was shown to be a critical factor in achieving superior performance. The competing Russian Tu-144 initially used a turbofan with reheat, but changed to a turbojet without reheat[4] with considerable improvement in performance.

Olympus-powered Concorde 216 (G-BOAF) on the final-ever Concorde landing, at Bristol, England

Development of the engine and engine accessories was the responsibility of Bristol Siddeley, while Snecma 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.

Increases in aircraft weight during the design phase led to a take-off thrust requirement which could not be met by the engine. The required shortfall of 20% was met with the introduction of partial reheat which was produced by SNECMA.[3]

The Olympus 593B was first run in November 1965. The B 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.[5] The B was dropped later from the designation. Snecma used an Olympus 301 in testing scaled models of the nozzle system.[6]

In June 1966, a complete Olympus 593 engine and variable geometry exhaust assembly was first run at Melun-Villaroche, Île-de-France, France. 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.[7]

In early 1966, the Olympus 593 produced 37,000 lb of thrust with reheat.[8]

In April 1967, the Olympus 593 ran for the first time in a high altitude chamber, at Saclay Île-de-France, France. 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 March 2, 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).

67 Olympus 593 engines were manufactured.[2]

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.[9] 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.[10]

Propulsion System Design[edit]

Engine[edit]

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.[11] 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)[3] 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.[12]

Intake[edit]

Concorde's intake system schematics
Concorde's intake system
G-AXDN, Duxford, close up of engines, with the scalloped thrust reversers prominent.

The Concorde's variable geometry intake, 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 pressure distribution (distortion) that can be tolerated by the compressor. Poor pressure recovery is an unacceptable loss for the intake compression process and unacceptable distortion causes engine surging (from loss of surge margin). If the engine is an afterburning turbojet the intake also has to supply cooling air for the hot afterburner duct and engine nozzle. Meeting all the above requirements over the relevant parts of the operating envelope was vital for Concorde to become a viable commercial aircraft. They were met with variable geometry and an intake control system that did not compromise the operation of the engine nor the control of the aircaft.

Supersonic pressure recovery is addressed by the number of shockwaves 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.[13] 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 centrebody to change the direction.[14] Subsonic pressure recovery is addressed by removal of the boundary layer (at the ramp bleed slot) and suitable shaping of the subsonic diffuser leading to the engine. The high pressure recovery for the Concorde intake at cruise gave an intake pressure ratio of 7.3:1.[15]

Shock waves gave rise to excessive boundary layer growth on the front ramp. The boundary layer was removed through the ramp bleed slot and bypassed the subsonic diffuser and engine where it would otherwise have caused excessive duct loss and unacceptable distortion at the engine.[16] 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.[16]

The dump doors were closed at cruise to prevent loss in thrust since air leaking from the duct does not contribute to the pressure recovery in the intake.[13]

At take-off, since the intake area was dimensioned for cruise, an auxiliary inlet was required to meet the higher engine flow. Distortion of the flow at the engine face also had to be addressed leading to an aerodynamic cascade with the auxiliary door.[16]

Forces from the internal airflow on the intake structure are rearwards (drag) on the initial converging section, where the supersonic deceleration takes place, and forwards on the diverging duct where subsonic deceleration takes place up to the engine entry. The sum of the 2 forces at cruise gave the 63% thrust contribution from the intake part of the propulsion system[12]

In order to achieve the necessary accuracy in the control of the intake ramp and spill positioning, it was found necessary to use a digital signal processor in the Air Intake Control Units. This was developed relatively late in the programme (~1972) but ensured the required fuel economy for transatlantic flights. The digital processor also accurately calculated the necessary engine speed scheduling to ensure an adequate surge margin under all engine and airframe operating conditions.

The intake control system had the unique ability to keep the powerplants operating correctly and to aid recovery, whatever the pilots, the aircraft and the atmosphere were doing in combination at the time.

The overall pressure ratio for the powerplant at Mach 2.0 cruise at 51,000 ft was about 82:1, with 7.3:1 from the intake and 11.3:1 from the 2 engine compressors.[15] The thermal efficiency with this high pressure ratio was about 43%.[3]

Exhaust nozzle[edit]

The variable geometry exhaust nozzle consisted of two "eyelids" which varied their position in the exhaust flow dependent on the flight regime, for example when fully closed (into the exhaust flow) they acted as thrust reversers, aiding deceleration from landing to taxi speed. In the fully open cruise position, together with engine nozzle, they formed an ejector nozzle to control the expansion of the exhaust. The eyelids formed the divergent passage while the engine exhaust ejected or pumped the secondary flow from the intake ramp bleed slot.

The expanding flow in the diverging section caused a forward thrust force on the exhaust nozzle, its 29% contribution to the overall propulsion system thrust at cruise.[12]

Variants[edit]

  • 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[edit]

Preserved examples of the Rolls-Royce/Snecma Olympus 593 are on display at the following museums:

The lobby at Henriksen Jet Center at Austin Executive Airport has an Olympus 593 Mk 610 engine on display.

Specifications (Olympus 593 Mk 610)[edit]

Turbine section of an Olympus 593 on display at the Fleet Air Arm Museum

Data from Jane's.[17][18]

General characteristics

  • Type: turbojet
  • Length: 4.039 m (13 ft 3 in)
  • Diameter: 1.212 m (47.75 in)
  • Dry weight: 3,175 kg (7,000 lb)

Components

  • Compressor: Axial flow, 7-stage low pressure, 7-stage high pressure
  • Combustors: Nickel alloy construction annular chamber, 16 vapourising burners, each with twin outlets
  • Turbine: High pressure single stage, low pressure single stage
  • Fuel type: Jet A1

Performance

See also[edit]

Related development
Comparable engines
Related lists

References[edit]

  1. ^ a b c "Olympus-the first forty years" Alan Baxter, RRHT No15, ISBN 1-9511710-9-7, p.135
  2. ^ a b [1]
  3. ^ a b c d e "Not Much of an Engineer" Sir Stanley Hooker An Autobiography, ISBN 1 85310 285, p.154
  4. ^ "Development of ABE Theory in Russia: Past, Present and Future" Ivanov, Central Institute of Aviation Motors, Moscow 111116
  5. ^ "Aero Engines", Flight, 6 January 1966: 28 
  6. ^ Flight April 1966
  7. ^ Testing of Concorde's engine on a Vulcan
  8. ^ "Historical Highlights", Flight International, 17 April 1969: 14 
  9. ^ http://www.flightglobal.com/pdfarchive/view/1974/1974%20-%200593.html
  10. ^ [2]
  11. ^ "Powerplant" concordesst.com
  12. ^ a b c "Brian Trubshaw Test Pilot" ISBN 0 7509 1838 1, Appendix VIIIb
  13. ^ a b "How Supersonic Inlets Work" J. Thomas Anderson, Copyright Lockheed Martin Corporation, published by Aero Engine Historical Society at "enginehistory.org"
  14. ^ Carlo Kopp (June 1997), Profile - The BAC TSR.2, Ausairpower.net, retrieved 19 February 2007, "First published in Australian Aviation," 
  15. ^ a b "Jet Propulsion" Nicholas Cumpsty, ISBN0 521 59674 2, p.149
  16. ^ a b c "Design and Development of an Air Intake for a Supersonic Transport Aircraft" Rettie and Lewis" Journal of Aircraft, November-December 1968 Vol. 5, No. 6
  17. ^ Janes - Online archive, Olympus 593 Retrieved: 9 November 2008
  18. ^ untitled

External links[edit]