Until regular commercial flights by Concorde ceased in October 2003, the Olympus turbojet was unique in aviation as the only turbojet with reheat powering a commercial aircraft.
The overall efficiency of the engine in supersonic cruising flight (supercruise) was about 43%, which at the time was the highest figure recorded for any normal thermodynamic machine.[3]
Development
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 1 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 thrust-specific fuel consumption (TSFC or simply SFC), engine pressure ratio, size and weight, 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.
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.[5] Britain was to have a larger share in production of the Olympus593 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.[5]
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]
In July 1964, two prototypes of the 593D engine ("D" for "Derivative", i.e., derived from the 22R) were produced. These two derivative engines were built to determine the validity of design concepts such as turbine stator and rotor cooling and testing the system at high environmental temperatures. They also demonstrated the need for larger engines, which were designated 593B.[6]
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.[7] The B was dropped later from the designation. Snecma used an Olympus 301 in testing scaled models of the nozzle system.[8]
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 (1,200km/h). During these tests, the 593 achieved 35,190lbf (157kN) thrust, which exceeded the specification for the engine.[9]
In early 1966, the Olympus 593 produced 37,000lbf (160kN) of thrust with reheat.[10]
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 (380km/h) after a ground run of 4,700 feet (1.4km).
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.[11] 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.[12]
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 120°C (393K; 248°F)[5] - the compressor drums and blades were made from titanium except for the last four HP stages, which were Nimonic 90[13] nickel alloy.[14] 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.[15]
Partial reheat providing a 20% thrust increase[3] was installed to give the take-off thrust required for Concorde to operate from existing runways, and for transonic acceleration from Mach 0.95 up to Mach 1.7; the aircraft flew supersonically without reheat above that speed. At cruise the engine's direct contribution (transferred by its mounts to the airframe) to forward thrust was 8% of propulsion system thrust. Its indirect contribution came from inducing air through the intake and pumping it (primary and secondary air) through the airframe nozzle: 63% derived from forward pressure against structures inside the air intake system, and 28% from forward pressure against the exhaust nozzles.[16][14]
All major components of the Olympus 593 were designed for a life of 25,000 hours, with the exception of the compressor and turbine blades, which were designed for a 10,000-hour life.[13] An engine installed on Concorde could be changed in 1 hour 50 minutes.[17]
Engine
Cutaway Olympus 593 showing LP compressor, HP compressor, combustor, turbines and reheat fuel ring and flame holder.
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.
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.
For comparison with the 593, a military full reheat requirement (50% thrust boost/3,000 degF[18]) needing three fuelling rings (hidden by red flameholders) for flame across the whole duct (GE J79 turbojet engine).
Intake
Concorde's intake system schematicsConcorde's intake system
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.[19] 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.[20] 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 pressure recovery achieved by Concorde's engine intakes at Mach 2 cruise gave an intake pressure ratio of 7.3:1.[21]
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.[22] 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.[22]
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.[19]
Since the intake area was optimal for cruise, an auxiliary inlet was provided to meet the higher engine air flow needed for take-off. Distortion of the flow at the engine face also had to be addressed, leading to an aerodynamic cascade with the auxiliary door.[22]
Forces from the internal airflow on the intake structure are rearward (drag) on the initial converging section, where the supersonic deceleration takes place, and forward on the diverging duct where subsonic deceleration takes place up to the engine entry. The sum of the two forces gave a 63% thrust contribution at cruise from the intake part of the propulsion system.[16]
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 in around 1972, relatively late in the programme, by the Electronics and Space Systems division of the British Aircraft Corporation at Filton. The Air Intake Control Units 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.
Concorde's Air Intake Control System also pioneered the use of digital data highways (multiplexedserial databuses) which connected the Air Intake Sensor Units that collected aerodynamic data at the nose of the aircraft (total pressure, static pressure, angle of attack and sideslip) and sent it to the Air Intake Control Units located nearer the air intakes, a distance of around 190ft (58m), using screened twisted pair cables to replace what would have been a much greater weight in aircraft wiring had only analogue signal wiring and pneumatic piping been used.
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 51,000ft (15,500m) was about 82:1, with 7.3:1 from the intake and 11.3:1 from the two engine compressors, far higher than any subsonic aircraft of the day, giving a correspondingly high overall efficiency of about 43%.[3][21]
Secondary nozzle
The first part of the exhaust system is the engine nozzle with adjustable area to control the engine, and known as the primary nozzle. The engine exhaust and the nacelle secondary air from the intake pass into the airframe-mounted combined modulating secondary nozzle and thrust reverser. This nozzle, developed by SNECMA, 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 the 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: a 29% contribution to the overall propulsion system thrust at cruise.[16]
During cruise at Mach 2.02 each Olympus 593 was producing around 10,000lbf (44kN) of thrust, equivalent to 36,000hp (27,000kW) per engine.[23]
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.[24]
Nozzle
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.
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.[25] They are fully closed to deflect the exhaust partially forward to produce reverse thrust.
Airframe secondary nozzle positions when A) taking off B) Mach 2 cruise C) thrust reversing.
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)[21]
A jet engine is a type of reaction engine, discharging a fast-moving jet of heated gas that generates thrust by jet propulsion. While this broad definition may include rocket, water jet, and hybrid propulsion, the term jet engine typically refers to an internal combustion air-breathing jet engine such as a turbojet, turbofan, ramjet, pulse jet, or scramjet. In general, jet engines are internal combustion engines.
A turbofan or fanjet is a type of airbreathing jet engine that is widely used in aircraft propulsion. The word "turbofan" is a combination of references to the preceding generation engine technology of the turbojet and the additional fan stage. It consists of a gas turbine engine which achieves mechanical energy from combustion, and a ducted fan that uses the mechanical energy from the gas turbine to force air rearwards. Thus, whereas all the air taken in by a turbojet passes through the combustion chamber and turbines, in a turbofan some of that air bypasses these components. A turbofan thus can be thought of as a turbojet being used to drive a ducted fan, with both of these contributing to the thrust.
The turbojet is an airbreathing jet engine which is typically used in aircraft. It consists of a gas turbine with a propelling nozzle. The gas turbine has an air inlet which includes inlet guide vanes, a compressor, a combustion chamber, and a turbine. The compressed air from the compressor is heated by burning fuel in the combustion chamber and then allowed to expand through the turbine. The turbine exhaust is then expanded in the propelling nozzle where it is accelerated to high speed to provide thrust. Two engineers, Frank Whittle in the United Kingdom and Hans von Ohain in Germany, developed the concept independently into practical engines during the late 1930s.
An afterburner is an additional combustion component used on some jet engines, mostly those on military supersonic aircraft. Its purpose is to increase thrust, usually for supersonic flight, takeoff, and combat. The afterburning process injects additional fuel into a combustor in the jet pipe behind the turbine, "reheating" the exhaust gas. Afterburning significantly increases thrust as an alternative to using a bigger engine with its attendant weight penalty, but at the cost of increased fuel consumption which limits its use to short periods. This aircraft application of "reheat" contrasts with the meaning and implementation of "reheat" applicable to gas turbines driving electrical generators and which reduces fuel consumption.
The Rolls-Royce Pegasus is a British turbofan engine originally designed by Bristol Siddeley. It was manufactured by Rolls-Royce plc. The engine is not only able to power a jet aircraft forward, but also to direct thrust downwards via swivelling nozzles. Lightly loaded aircraft equipped with this engine can manoeuvre like a helicopter. In particular, they can perform vertical takeoffs and landings. In US service, the engine is designated F402.
The Rolls-Royce Olympus was the world's second two-spool axial-flow turbojet aircraft engine design, first run in May 1950 and preceded only by the Pratt & Whitney J57, first-run in January 1950. It is best known as the powerplant of the Avro Vulcan and later models in the Concorde SST.
The Pratt & Whitney J58 is an American jet engine that powered the Lockheed A-12, and subsequently the YF-12 and the SR-71 aircraft. It was an afterburning turbojet engine with a unique compressor bleed to the afterburner that gave increased thrust at high speeds. Because of the wide speed range of the aircraft, the engine needed two modes of operation to take it from stationary on the ground to 2,000 mph (3,200 km/h) at altitude. It was a conventional afterburning turbojet for take-off and acceleration to Mach 2 and then used permanent compressor bleed to the afterburner above Mach 2. The way the engine worked at cruise led it to be described as "acting like a turboramjet". It has also been described as a turboramjet based on incorrect statements describing the turbomachinery as being completely bypassed.
A propelling nozzle is a nozzle that converts the internal energy of a working gas into propulsive force; it is the nozzle, which forms a jet, that separates a gas turbine, or gas generator, from a jet engine.
The Armstrong Siddeley Sapphire is a British turbojet engine that was produced by Armstrong Siddeley in the 1950s. It was the ultimate development of work that had started as the Metrovick F.2 in 1940, evolving into an advanced axial flow design with an annular combustion chamber that developed over 11,000 lbf (49 kN). It powered early versions of the Hawker Hunter and Handley Page Victor, and every Gloster Javelin. Production was also started under licence in the United States by Wright Aeronautical as the J65, powering a number of US designs. The Sapphire's primary competitor was the Rolls-Royce Avon.
The Bristol Siddeley Orpheus is a single-spool turbojet developed by Bristol Siddeley for various light fighter/trainer applications such as the Folland Gnat and the Fiat G.91. Later, the Orpheus formed the core of the first Bristol Pegasus vectored thrust turbofan used in the Harrier family.
The Rolls-Royce/SNECMA M45H is an Anglo-French medium bypass ratio turbofan produced specifically for the twin-engined VFW-Fokker 614 aircraft in the early 1970s.
The Bristol Siddeley BS.100 is a British twin-spool, vectored thrust, turbofan aero engine that first ran in 1960. The engine was designed and built in limited numbers by Bristol Siddeley Engines Limited. The project was cancelled in early 1965.
A variable cycle engine (VCE), also referred to as adaptive cycle engine (ACE), is an aircraft jet engine that is designed to operate efficiently under mixed flight conditions, such as subsonic, transonic and supersonic.
The General Electric GE4 turbojet engine was designed in the late 1960s as the powerplant for the Boeing 2707 supersonic transport. The GE4 was a nine-stage, single-shaft, axial-flow turbojet based largely on the General Electric YJ93 which powered the North American XB-70 bomber. The GE4 was the most powerful engine of its era, producing 50,000 lbf (220 kN) dry, and 65,000 lbf (290 kN) with afterburner. The Boeing 2707 was cancelled in 1971, putting an end to further work on the GE4.
This article briefly describes the components and systems found in jet engines.
The Volvo RM8 is a low-bypass afterburning turbofan jet engine developed for the Saab 37 Viggen fighter. An augmented bypass engine was required to give both better fuel consumption at cruise speeds and higher thrust boosting for its short take-off requirement than would be possible using a turbojet. In 1962, the civil Pratt & Whitney JT8D engine, as used for airliners such as the Boeing 727, was chosen as the only engine available which could be modified to meet the Viggen requirements. The RM8 was a licensed-built version of the JT8D, but extensively modified for supersonic speeds, with a Swedish-designed afterburner, and was produced by Svenska Flygmotor.
An airbreathing jet engine is a jet engine in which the exhaust gas which supplies jet propulsion is atmospheric air, which is taken in, compressed, heated, and expanded back to atmospheric pressure through a propelling nozzle. Compression may be provided by a gas turbine, as in the original turbojet and newer turbofan, or arise solely from the ram pressure of the vehicle's velocity, as with the ramjet and pulsejet.
The familiar study of jet aircraft treats jet thrust with a "black box" description which only looks at what goes into the jet engine, air and fuel, and what comes out, exhaust gas and an unbalanced force. This force, called thrust, is the sum of the momentum difference between entry and exit and any unbalanced pressure force between entry and exit, as explained in "Thrust calculation".
The Rolls-Royce Olympus turbojet engine was developed extensively throughout its production run, the many variants can be described as belonging to four main groups.
The Boom Symphony is a medium-bypass turbofan engine under development by Boom Technology for use on its Overture supersonic airliner. The engine is designed to produce 35,000 pounds of thrust at takeoff, sustain Overture supercruise at Mach 1.7, and burn sustainable aviation fuel exclusively.
1 2 "Concorde technical specs: powerplant". Concorde SST. Gordon Roxburgh. Archived from the original on 29 October 2023. Retrieved 29 October 2023. During the Supersonic cruse only 8% of the power is derived by the engine with the other 29% being from Nozzles and an impressive 63% from the intakes.
1 2 "How Supersonic Inlets Work" J. Thomas Anderson, Copyright Lockheed Martin Corporation, published by Aero Engine Historical Society at "enginehistory.org"
↑ Carlo Kopp (June 1997), "Profile - The BAC TSR.2", Australian Aviation, 1997 (June), Ausairpower.net, retrieved 19 February 2007, First published in Australian Aviation,
1 2 3 "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
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