Exhaust Gas Stream Velocity
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Exhaust Gas Stream Velocity
Dear Reader,
My Dad was asking me if I knew at what speed the stream of exhaust gases is doing when it comes out the back of a jet engine. He used the example of a B747 at 400 kts. I appreciate the answer my depend on the type of engine concerned, power setting, atmospheric conditions and so on. I was unable to answer his question and I was wondering if there were any RR, P&W, or GE types out there who could provide an answer. Are there any takers?
My Dad was asking me if I knew at what speed the stream of exhaust gases is doing when it comes out the back of a jet engine. He used the example of a B747 at 400 kts. I appreciate the answer my depend on the type of engine concerned, power setting, atmospheric conditions and so on. I was unable to answer his question and I was wondering if there were any RR, P&W, or GE types out there who could provide an answer. Are there any takers?
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Good question.
I believe on civil engines, the core exit velocity (the hot bit) is normally just below the speed of sound. At a temperature of, say, 700K (about 400 deg.C) the speed of sound would be about 517 m/s (just over 1000 kts) I am a bit rusty with the figures, and this sounds a little high, but I would say at least about 800 kts. The speed out of the bypass duct would be considerably lower than this.
Re-heated military engines can have higher exit velocities, because they use convergent-divergent nozzles to accelerate the flow to supersonic speeds, but this is very wasteful of fuel.
Please, anybody, correct me if my theory is a little rusty.
Hope this helps.
Pie
I believe on civil engines, the core exit velocity (the hot bit) is normally just below the speed of sound. At a temperature of, say, 700K (about 400 deg.C) the speed of sound would be about 517 m/s (just over 1000 kts) I am a bit rusty with the figures, and this sounds a little high, but I would say at least about 800 kts. The speed out of the bypass duct would be considerably lower than this.
Re-heated military engines can have higher exit velocities, because they use convergent-divergent nozzles to accelerate the flow to supersonic speeds, but this is very wasteful of fuel.
Please, anybody, correct me if my theory is a little rusty.
Hope this helps.
Pie
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Sounds reasonable. I would hesitate to offer specific velocities without specific data .. but, then again, that might be just because I am a boring old conservative engineer.
Re 1000kt, making the usual assumptions, the normal sonic calculation is
a1/a2 = square root(T1/T2)
with the temperatures measured in degrees absolute.
For the example you quoted
a1 = 661 x square root(700/288)
= 1030 kt
or thereabouts ....
Re 1000kt, making the usual assumptions, the normal sonic calculation is
a1/a2 = square root(T1/T2)
with the temperatures measured in degrees absolute.
For the example you quoted
a1 = 661 x square root(700/288)
= 1030 kt
or thereabouts ....
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John
As always, I was using:
a = Sqrt(gamma x R x T)
At a guess, gamma would be between 1.3 and 1.35 for a core nozzle outlet mixture, and I would be the first to admit that the temperature is pure guesswork. I am glad you were able to corroborate the principle, though.
Thanks.
Pie
As always, I was using:
a = Sqrt(gamma x R x T)
At a guess, gamma would be between 1.3 and 1.35 for a core nozzle outlet mixture, and I would be the first to admit that the temperature is pure guesswork. I am glad you were able to corroborate the principle, though.
Thanks.
Pie
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The RB211-524s at take-off power had jet velocities around 1500 ft/s in the hot jet and about 1000 ft/s in the bypass, which as previously stated is locally subsonic.
More recent civil aircraft will have lower velocities, mainly because of noise regulations. I believe the Trent is around 1100-1200 ft/s but can't be sure.
More recent civil aircraft will have lower velocities, mainly because of noise regulations. I believe the Trent is around 1100-1200 ft/s but can't be sure.
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Thanks everybody for your responses. I'm sure that will satisfy my Dad's query. By the way, he was also wondering approximately how much of the total (if any of the total) forward thrust of the engine would be provided by the suction action of the fan and compressors drawing air in. That's way beyond my understanding of jet engines. Ideas? Thanks again.
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It is true that a large proportion of thrust is produced by a pressure rise across the fan, and when the aircraft is stationary, I suppose there must be a static pressure reduction at the front surface of the fan blades, since the fan is downstream of the intake flow (clearly!)
Again correct me if I'm wrong, but:
p0 = P0 - 1/2(rho*V^2)
and
p1 = P0 + Delta P
where p is static pressure, P is total pressure, stations 0 and 1 are immediately before and after the fan, V is the air velocity through the fan, and and Delta P is the pressure loss through the bypass duct. This assumes the engine is at rest and also assumes incompressible flow.
Pressure difference would be p1 - p0. Multiply by annular fan area to get the thrust from the fan.
So you could say that the thrust from the fan is down to the static pressure reduction at the front of the engine, but I would prefer to break the thrust down into pressure differences across individual engine components, or simply consider the momentum balance across the whole engine. These are far more rhobust ways of explaining where thrust comes from.
I would recommend you get hold of a copy of "The Jet Engine" - Rolls Royce - ISBN 0 902121 2 35
Pie
Again correct me if I'm wrong, but:
p0 = P0 - 1/2(rho*V^2)
and
p1 = P0 + Delta P
where p is static pressure, P is total pressure, stations 0 and 1 are immediately before and after the fan, V is the air velocity through the fan, and and Delta P is the pressure loss through the bypass duct. This assumes the engine is at rest and also assumes incompressible flow.
Pressure difference would be p1 - p0. Multiply by annular fan area to get the thrust from the fan.
So you could say that the thrust from the fan is down to the static pressure reduction at the front of the engine, but I would prefer to break the thrust down into pressure differences across individual engine components, or simply consider the momentum balance across the whole engine. These are far more rhobust ways of explaining where thrust comes from.
I would recommend you get hold of a copy of "The Jet Engine" - Rolls Royce - ISBN 0 902121 2 35
Pie
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A large proportion of the engine's thrust comes from pressure differentials across the sidewall structures, especially in the inlet nacelle area. The idea of "suction" across the fan is not really valid, in the same way as for propellers.
The following thread in this forum has some interesting data for a pointy aircraft which goes quite quickly...
"Concorde's engines at Mach 2"
The following thread in this forum has some interesting data for a pointy aircraft which goes quite quickly...
"Concorde's engines at Mach 2"
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Although statically you could concieve of the fan 'sucking in' air, in flight their is a pressure rise in the diffuser section of the intake as the air slows down before going through the fan, statically the air has to be accelerated first.
This creates a momentum drag component on the engine due to the force required to slow the air down. Of course a large amount of the power required to overcome this is recovered as the air expands again through the nozzles.
This creates a momentum drag component on the engine due to the force required to slow the air down. Of course a large amount of the power required to overcome this is recovered as the air expands again through the nozzles.
