Olympus 593 Engines

Olympus 593
Preserved Olympus 593 engine at the Imperial War Museum Duxford
TypeTurbojet
National originUnited Kingdom/France
ManufacturerRolls-Royce Limited/Snecma
First runJune 1966
Major applicationsConcorde
Number built67
Developed fromRolls-Royce Olympus

To be economically viable, Concorde needed to be able to fly reasonably long distances, and this required high efficiency. For optimum supersonic flight, turbofan engines were considered, but rejected, as due to their large master cross-section they would cause excessive drag. Turbojets were found to be the best choice of engines. The quieter high bypass turbofan engines such as used on Boeing 747s could not be used. The engine chosen was the twin spool Rolls-Royce/Snecma Olympus 593, a version of the Olympus originally developed for the Vulcan bomber, developed into an afterburning supersonic engine for the BAC TSR-2 strike bomber and then adapted for Concorde.

The inlet design for Concorde's engines was critical. All conventional jet engines can take in air at only around Mach 0.5; therefore the air needs to be slowed from the Mach 2.0 airspeed that enters the engine inlet. In particular, Concorde needed to control the shock waves that this reduction in speed generates to avoid damage to the engines. This was done by a pair of intake ramps and an auxiliary flap, whose position was moved during flight to slow the air down. The ramps were at the top of the engine compartment and moved down and the auxiliary flap moved both up and down allowing air to flow in or out. During takeoff, when the engine's air demand was high, the ramps were flat at the top and the auxiliary flap was in, allowing more air to enter the engine. As the aircraft approached Mach 0.7, the flap closed; at Mach 1.3, the ramps came into effect, removing air from the engines which was then used in the pressurisation of the cabin. At Mach 2.0, the ramps had covered half their total possible distance. They also helped reduce the work done by the compressors as they not only compressed the air but also increased the air temperature.

Engine failure causes large problems on conventional subsonic aircraft; not only does the aircraft lose thrust on that side but the engine is a large source of drag, causing the aircraft to yaw and bank in the direction of the failed engine. If this had happened to Concorde at supersonic speeds, it could theoretically have caused a catastrophic failure of the airframe. However, during an engine failure air intake needs are virtually zero, so in Concorde the immediate effects of the engine failure were countered by the opening of the auxiliary flap and the full extension of the ramps, which deflected the air downwards past the engine, gaining lift and streamlining the engine, minimising the drag effects of the failed engine. In tests, Concorde was able to shut down both engines on the same side of the aircraft at Mach 2 without any control problems.

The aircraft used reheat (afterburners) at take-off and to pass through the transonic regime (i.e. "go supersonic") between Mach 0.95 and Mach 1.7, and were switched-off at all other times. The engines were capable of reaching Mach 2 without reheat, but it was discovered that it burnt more fuel that way, since the aircraft spent much longer flying in the high-drag transonic regime even though reheat is inefficient.

Due to jet engines being highly inefficient at low speeds, Concorde burned two tonnes of fuel taxiing to the runway. To conserve fuel only the two outer engines were run after landing. The thrust from two engines was sufficient for taxiing to the ramp due to low aircraft weight upon landing at its destination.

However, when operating Concorde at its design point at Mach 2, it was the world's most efficient jet engine.

Concorde's ramp system schematics

The Rolls-Royce/Snecma Olympus 593 was a reheated (afterburning) turbojet which powered the supersonic airliner Concorde. Rolls-Royce Limited and Snecma developed the engine from the Rolls-Royce Olympus. Until Concorde's regular commercial flights ceased the Olympus turbojet was unique in aviation as the only afterburning turbojet powering a commercial aircraft.

Design and development

The Olympus 593 project was started in 1964, using the BAC TSR-2's Olympus 320 as a basis for development. Bristol Siddeley of the UK and Snecma Moteurs of France were to share the project. Acquiring Bristol Siddeley in 1966, Rolls-Royce continued as the British partner. The early stages validated the basic design concept but many studies were required to achieve desired specifications, e.g.

  • The critical factor fuel consumption
  • Pressure Ratio
  • Weight/Size
  • Turbine entry temperature

Initially, engineers studied using turbojets or 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, but quickly changed to a turbojet with considerable improvement in performance.

Rolls-Royce carried out the development of the original Bristol Siddeley Olympus and engine accessories, while Snecma was responsible for the variable engine inlet system, the exhaust nozzle/thrust reverser, the afterburner and the noise attenuation system. Britain was to have a larger share in production of the Olympus 593 as France had a larger share in fuselage production.

In June 1966 a complete Olympus 593 engine and variable geometry exhaust assembly was first run at Melun-Villaroche, le-de-France, France. At Bristol, flight tests began using a RAF Vulcan bomber with the engine attached to its underside. 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 requirements of the engine.

Concorde's ramp system

In April 1967 the Olympus 593 ran for the first time in a high altitude chamber, at Saclay le-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.

At 15:40 on the 2nd March 1969 Concorde prototype 001, captained by chief test pilot Andre Turcat, started its first take off run, with afterburners lit. The four Olympus 593 engines accelerated the aircraft, and after 4,700 feet (1.4 km) of runway and at a speed of 205 knots (380 km/h), captain Turcat lifted the aircraft off for the first time.

67 engines in the end were manufactured.

Plans were drawn up for a quieter and more powerful version of the engine with an extra turbine section and a larger-diameter air compressor that would have eschewed the reheat and added sound-deadening; this would have improved efficiency across the board and permitted rather greater range and opened up new routes, particularly across the Pacific as well as transcontinental across America. However, the poor sales of Concorde meant that this plan for a Concorde 'B' was never put into practice.

Technical description

A jet engine draws air in at the front and compresses it. The air then combines with fuel and the engine burns the resulting mixture. The combustion greatly increases the volume of the gases which are then exhausted out of the rear of the engine. The Olympus engine took this gas jet and passed it through straightening vanes - to obtain a laminar flow. This gas jet then entered the afterburning jet pipe where a ring of fuel injectors sprayed fuel onto the hot exhaust gases. The resulting combustion greatly improved thrust, although it also lead to high fuel consumption.

The afterburning section was longer than the engine itself (as was the case with all early turbojets) but the thrust of the Olympus 302 rose to 30,610 lbf (136 kN).

Inlet system

Although not strictly considered part of the engine, the variable engine inlet system was vital to the Olympus 593 on Concorde as supersonic airflow at the engine face could create shockwaves that would lead to engine surge and failure. The intake featured variable ramps which altered the intake area which slowed the intake air from supersonic to subsonic speed. This was achieved by positioning the ramps such that shockwaves were created at the inlet, air passing through these shockwaves was slowed. The inlet air was further decelerated as the intake area increased closer to the engine (speed of air flow decreases as area increases.) Any excess air was expelled through doors on the underside of the nacelle and some intake air was bypassed around the engine and mixed with the exhaust - to increase thrust and keep the engine cool. The TSR-2 had used a semicircular design with the shock front adjusted by a translating centrebody to overcome the same problem.

The overall pressure ratio of the air by the inlet system as well as the compression within the engine was as high as 80:1 at supersonic speeds. This gave the engine world-beating efficiency.

G-AXDN, Duxford, close up of engines, with the scalloped thrust reversers prominent.

Exhaust nozzle

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.

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

Specifications (Olympus 593 Mk 610)

General characteristics

  • Type: Turbojet
  • Length: 4039 mm (159 in)
  • Diameter: 1212 mm (47.75 in)
  • Dry weight: 3175 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

  • Maximum Thrust: wet: 169.2 kN (38,050 lbf) dry: 139.4 kN (31,350 lbf)
  • Overall pressure ratio: 15.5:1
  • Specific fuel consumption: 1.195 (cruise), 1.39 (SL) lb/(hlbf)
  • Thrust-to-weight ratio: 5.4

Control system

  • World's first FADEC control system

Jetpipe

  • Straight pipe with pneumatically operated convergent nozzle
  • Single ring afterburner
  • 'Eyelids' which act as variable divergent nozzles/thrust reversers