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Rolls Royce SNECMA Olympus 593 Mrk 610 Turbojet

The Rolls Royce Olympus Turbojet is somewhat of a unique engine in that it is the only afterburning turbojet engine used to power a commercial aircraft. The engine is a development of the Bristol-Siddeley Olympus that was used to power the British Avro Vulcan long range bomber. Rolls-Royce carried out the development of the Bristol Siddeley powerplant for the Concorde, while Snecma designed the variable engine inlet system, the afterburner, and the variable geometry exhaust nozzle/thrust reverser. In it's final iteration, the Olympus 593 could produce a maximum thrust output of 38,050 lbs of thrust, with afterburner. The engine produces about 32,000 lbs of thrust at full power without afterburners.

The Rolls Royce Olympus 593 is a fairly complex turbojet engine. It is a two spool engine, featuring a split compressor for increased engine efficiency and power, as well as improved acceleration characteristics. Air enters the engine face and passes through a fixed 5 spoke zero-swirl inlet guide vane ring. The air is drawn in and compressed by the 7 stage axial low pressure compressor. The low pressure compressor is driven by a single stage turbine. After the low pressure compressor, the compressed air is fed to the high pressure compressor, which is also a 7 stage axial compressor, and is also driven by a single stage turbine. The high pressure shaft spins coaxially with, and on the outside of the low pressure shaft. The two shafts rotate independently of one another. From the high pressure compressor, air is fed to the diffuser and then to the annular through-flow combustor. Fuel is sprayed into the combustor via 16 duplex vaporizing burners, where it mixes with the compressed air and ignites to provide high energy gas to drive the turbines and provide the propulsive jet. The gas is accelerated through the first stage nozzle, which is cooled with compressor bleed air, and is then expanded through the single stage axial high pressure turbine, also cooled. The high pressure turbine drives the high pressure compressor and the accessory gearbox via a tower shaft and bevel gear in front of the high pressure compressor. The high pressure spool rides on the outside of the low pressure shaft, coaxially. The gas is then accelerated through the second stage turbine nozzle, and is expanded through the single stage axial low pressure turbine, which is also cooled with bleed air.

The remaining energy in the gas can then be used to provide thrust to propel the aircraft. The gas passes a set of straightening vanes which remove any swirl, and then enters the afterburning jet pipe. In the front of the jet pipe is a single ring of fuel injectors which spray fuel into the pipe when the afterburner is activated. The jet pipe acts as a combustion chamber for the afterburner, and allows for stable combustion of the fuel. Afterburning increases the energy of the exhaust stream dramatically, but also results in an even more dramatic increase in fuel consumption. Use of the afterburner is therefore limited to takeoff and supersonic acceleration. A fixed convergent nozzle at the end of the jetpipe creates a throat that accelerates the airflow to high velocity to provide thrust. The secondary jet nozzle is a variable area divergent nozzle, with two "eyelids" that vary their position to tailor the jet air flow to match power setting and flight regime. The nozzles are also varied to control N2 (low pressure spool) rpm, independent of the high pressure spool speed. The nozzle is computer controlled. When the eyelids are fully closed, they act as thrust reversers which aid in the deceleration from landing speed to taxi speed.

The Olympus engines on the Concorde would not function without the complex intake system located in the front of the Concorde engine nacelles. The intake system features variable ramps which alter the intake area. At low aircaft speed, the ramps are fully retracted, to allow the maximum area for the engine to draw in air. However, supersonic airflow at the engine face would create shockwaves that would cause engine surge and failure. The airflow must be slowed from supersonic speed to subsonic before it enters the engine. This is accomplished by adjusting the position of the ramps. This creates oblique shockwaves at the engine inlet causing the airflow to slow down as it passes through the shocks. The air is decelerated further as the intake area widens in front of the engine face. Excess air pressure is released through spill doors on the underside of the nacelle. Some intake air is bypassed around the engine and mixed with the exhaust, to keep the engine cool and to provide additional thrust.

Engine control, as well as control of the intake ramps and the exhaust nozzles is via computer control. N1 rpm is a direct function of fuel flow to the combustor, while N2 rpm is a function of fuel flow as well as exhaust nozzle position. When the pilot sets takeoff power, fuel flow is increased to bring N1 up to 100%, and then the nozzle is automatically adjusted to maintain 100% N2. Engine accessories are driven off of the two accessory gearboxes, which are driven off of the high pressure spool, and are mounted on the underside of the engine. Accessories include the oil pressure and scavenge pumps, the fuel pump, and the fuel controller unit, as well as an aircraft generator and hydraulic pumps. Engine starting is accomplished by a pneumatic air turbine starter motor, mounted on the accessory gearbox. Engine ignition is via high energy electric spark.

Rolls-Royce/SNECMA Olympus 593 Mrk 610 Afterburning Turbojet

  • Type: Dual spool axial flow turbojet with afterburning
  • Inlet: Axial
  • Compressor: Split, twin spool. 7 stage axial LPC, 7 stage axial HPC
  • Burner: Annular, 16 fuel nozzles
  • Turbine: Dual spool, single stage axial high pressure turbine, single stage axial low pressure turbine
  • Exhaust: Constant diameter afterburning jet pipe, convergent primary nozzle, variable area divergent secondary nozzle with thrust reverser.
  • Thrust Rating: 38,050 lbs. of thrust with afterburner
  • Maximum Dry Thrust: 31,500 lbs. of thrust
  • Weight: 7,000 lbs.
  • Thrust/weight: 5.4:1
  • Air mass flow: 410 lbs/sec
  • Compression Ratio: 15.5:1
  • Specific Fuel Consumption: .59 lb/lbt/hr (dry), 1.13 lb/lbt/hr (with afterburner)
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