Concorde TIT at Mach 2
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Concorde TIT at Mach 2
I would appreciate any help regarding any information someone might have for the TIT of the Concorde Olympus at Mach 2 - thanks in advance!
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There is a cracking good thread on Concorde here in Tech Log with generous contributions (2000+ posts) from designers, pilots, and mechanics that would be an ideal place to put your query, JCA.
PPRunE - Tech Log - Concorde Question
PPRunE - Tech Log - Concorde Question
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Flightglobal archive from 1971 has some information.
http://https://www.flightglob...0-%200615.html
"The Olympus is an essentially simple two-spool afterburning turbojet located behind a sophisticated variablegeometry intake, designed and built by BAC, which has overall powerplant responsibility and builds the nacelles. The engine has to accept intake temperatures of more than 120°C during Mach 2 cruise, so extensive use of titanium is made at the front end of the compressor for better creep resistance."
http://https://www.flightglob...0-%200615.html
"The Olympus is an essentially simple two-spool afterburning turbojet located behind a sophisticated variablegeometry intake, designed and built by BAC, which has overall powerplant responsibility and builds the nacelles. The engine has to accept intake temperatures of more than 120°C during Mach 2 cruise, so extensive use of titanium is made at the front end of the compressor for better creep resistance."
That is compressor inlet temp, GR. And indeed the Tmo was 127°C, on the nose. After compression of 82:1 (from both the intake ram/shock and the spinning compressors themselves) the Turbine Inlet/Entry Temp was much higher.
I found the following from another Flight Global article on the development of the Olympus 593 from 1969 (Author: M. H. Beanland, Asst. Chief Development Engineer, Olympus 593, Rolls-Royce Bristol Engine Division): https://www.flightglobal.com/FlightP...20-%201603.PDF
"The need for this** can be appreciated when it is realised that the temperature of the air leaving the compressors at cruise, and before any fuel has been burnt in it, exceeds 600°C; and that although the cruise turbine-entry temperature is substantially lower than that at take-off, the actual turbine-blade metal temperature is higher at cruise than at take-off, due to the much higher cooling-air temperature."
** "this" was an intake heater to simulate supersonic cruise conditions when testing the 593 on a Vulcan at subsonic speeds
I found the following from another Flight Global article on the development of the Olympus 593 from 1969 (Author: M. H. Beanland, Asst. Chief Development Engineer, Olympus 593, Rolls-Royce Bristol Engine Division): https://www.flightglobal.com/FlightP...20-%201603.PDF
"The need for this** can be appreciated when it is realised that the temperature of the air leaving the compressors at cruise, and before any fuel has been burnt in it, exceeds 600°C; and that although the cruise turbine-entry temperature is substantially lower than that at take-off, the actual turbine-blade metal temperature is higher at cruise than at take-off, due to the much higher cooling-air temperature."
** "this" was an intake heater to simulate supersonic cruise conditions when testing the 593 on a Vulcan at subsonic speeds
Just curious
why would there be anything unique about a Turbine Inlet Temp for a Concorde compare to other installations of the day?
I would think that the actual metal temp is a heat transfer thingie between the TIT out of the burner vs the cooling air temp from wherever you can tap enough pressure to flow to the turbine.
Is it not possible to move farther forward in the Concorde to get pressures and temperature for cooling the turbine in line with other installations of the same type engines?
why would there be anything unique about a Turbine Inlet Temp for a Concorde compare to other installations of the day?
I would think that the actual metal temp is a heat transfer thingie between the TIT out of the burner vs the cooling air temp from wherever you can tap enough pressure to flow to the turbine.
Is it not possible to move farther forward in the Concorde to get pressures and temperature for cooling the turbine in line with other installations of the same type engines?
I don't know much about the Olympus engine, but I read an article quite a while ago on the still-born Boeing SST that claimed it's engine was going to be the first commercial engine to use 'film cooled' first stage turbine blades (that's where the cooling air exits along the surface of the blade and creates a 'film' of cooling air over the blade - common practice in current turbine engines).
No, I can't vouch for it's accuracy
No, I can't vouch for it's accuracy
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why would there be anything unique about a Turbine Inlet Temp for a Concorde compare to other installations of the day?
I would think that the actual metal temp is a heat transfer thingie between the TIT out of the burner vs the cooling air temp from wherever you can tap enough pressure to flow to the turbine.
Is it not possible to move farther forward in the Concorde to get pressures and temperature for cooling the turbine in line with other installations of the same type engines?
Can't see any reason in principle why one should not tap cooling air earlier in the cycle, but I was under the impression that the Ol 593 cooling technology was state of the art for a 1960's engine
Last edited by CliveL; 17th May 2018 at 13:20. Reason: Re-read Mike Beanland's comments !
Can't see any reason in principle why one should not tap cooling air earlier in the cycle, but I was under the impression that the Ol 593 cooling technology was state of the art for a 1960's engine
It was probably just simpler to take the engine "off-the-shelf" and be done with it as far as cooling air to the turbine.
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I can confirm Clive's assessment:
The HP turbine cooling air came from the 5th stage of the 7 stage HP compressor. Incidently 5th stage air required for cooling and sealing the compressor bearings needed a heat exchanger. ref SAE 670316 Fig.11. So it seems the lowest pressure source possible was used for the job and no heat exchanger was needed either.
I read somewhere that there was no compromising the compressor design for the M2 cruise condition so as well as the definition of blade angles,etc. that would include minimizing the air taken from the gas path to the internal (secondary) air system for cooling,etc.
See https://en.wikipedia.org/wiki/Jet_en...e#Introduction for non-technical explanations on sources of waste (entropy) generation in jet engines including secondary air system. This article is a recent revamp of the previous which was considered too technical.
The HP turbine cooling air came from the 5th stage of the 7 stage HP compressor. Incidently 5th stage air required for cooling and sealing the compressor bearings needed a heat exchanger. ref SAE 670316 Fig.11. So it seems the lowest pressure source possible was used for the job and no heat exchanger was needed either.
I read somewhere that there was no compromising the compressor design for the M2 cruise condition so as well as the definition of blade angles,etc. that would include minimizing the air taken from the gas path to the internal (secondary) air system for cooling,etc.
See https://en.wikipedia.org/wiki/Jet_en...e#Introduction for non-technical explanations on sources of waste (entropy) generation in jet engines including secondary air system. This article is a recent revamp of the previous which was considered too technical.
