Spool systems in compressors

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Main Question or Discussion Point

why the spools(twin spool,triple spool) are used in compressors,why the stages in compressor must be increased
 
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Answers and Replies

  • #2
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I think a very basic google search on jet engines would provide you with this answer.
 
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  • #3
FredGarvin
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Your question doesn't make a whole lot of sense to me in the current form Can you restate it? It appears that you are asking why if you have multiple spool engines must the number of compressor stages increase. Is this correct?
 
  • #4
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Your question doesn't make a whole lot of sense to me in the current form Can you restate it? It appears that you are asking why if you have multiple spool engines must the number of compressor stages increase. Is this correct?
i mean y do we need them
 
  • #5
minger
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Rather than just have say one 20:1 pressure ratio stage?
 
  • #6
FredGarvin
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Compressors, especially axial compressors are very limited as to the pressure ratio across them allowed. That means more stages required for a higher required pressure ratio.
 
  • #7
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By "allowed," I take it to mean you're referring to what the laws of physics allow? Or is it simply a case design optimization, whereby obtaining the same, final compression ratio via multi-stage axial compression does not require planetary gearing or substantially greater weight or strength of materials than would a design with fewer stages?
 
  • #8
FredGarvin
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By "allowed," I take it to mean you're referring to what the laws of physics allow? Or is it simply a case design optimization, whereby obtaining the same, final compression ratio via multi-stage axial compression does not require planetary gearing or substantially greater weight or strength of materials than would a design with fewer stages?
I mean allowed in that compressors are constantly working against an adverse pressure gradient which places a lot of limitations on what kind of delta P across a stage you can have. Has anyone ever noticed that all jet engines (with axial compressors) have a much larger number of compressor stages than the turbine?
 
  • #9
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I mean allowed in that compressors are constantly working against an adverse pressure gradient which places a lot of limitations on what kind of delta P across a stage you can have.
You can design for a very high stage differential pressure, if you want. It's just terribly inefficient. :)

Has anyone ever noticed that all jet engines (with axial compressors) have a much larger number of compressor stages than the turbine?
Yes. Why is that, Fred?
 
  • #10
FredGarvin
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You can design for a very high stage differential pressure, if you want. It's just terribly inefficient. :)
Not in an axial machine. You can with a radial/centrif if you want.



Yes. Why is that, Fred?
One word answer: separation.
 
  • #11
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Not in an axial machine. You can with a radial/centrif if you want.
Yes, http://en.wikipedia.org/wiki/Axial_compressor#Spools". As I said, they're simply much less efficient.

One word answer: separation.
Separation occurs when turbines designed for a one pressure differential encounter a significantly greater pressure differential. It's a consequence of the design, as efficiency is highest when operated closest to the stall line.

Turbines with fewer stages and specifically designed for higher per stage pressure differentials do not experience separation within their design parameters. However, as I've mentioned, they're significantly less efficient than turbines designed for optimal efficiency, and which, as a consequence of their design, have more stages.

In short, it's entirely doable. It's simply not practical.
 
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  • #12
FredGarvin
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Separation occurs when turbines designed for a one pressure differential encounter a significantly greater pressure differential. It's a consequence of the design, as efficiency is highest when operated closest to the stall line.

Turbines with fewer stages and specifically designed for higher per stage pressure differentials do not experience separation within their design parameters. However, as I've mentioned, they're significantly less efficient than turbines designed for optimal efficiency, and which, as a consequence of their design, have more stages.

In short, it's entirely doable. It's simply not practical.
You're right. What do I know.
 
  • #13
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As an engine R&D guy, you know a great deal, Fred. This was simply a matter of extrapolating from existing design, in which case you were correct, vs considering a possible, but far less efficient, and therefore impractical design, for the concept mentioned by anvesh111 and minger.
 
  • #14
minger
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Whoa whoa, apparently your sarcasto-meter didn't pick me up properly.

We do centrifugal compressors, but I know enough about axial systems to know that a 20:1 pressure ratio (for example) on a 5' fan blade is going to cause some issues.
 
  • #15
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Whoa whoa, apparently your sarcasto-meter didn't pick me up properly.

We do centrifugal compressors, but I know enough about axial systems to know that a 20:1 pressure ratio (for example) on a 5' fan blade is going to cause some issues.
Absolutely. As I said, it's not impossible. Merely grossly inefficient, and therefore impractical.

A similar, and familiar approach would involve a ramjet. Below Mach 0.5, they have almost zero thrust due to poor compression ratios, and below 600 kts, they're grossly inefficient. Between Mach 2 and 4, however, they'll outperform any turbojet.

The issue is one of compression. In a turbojet, we obtain efficiency by using multiple compression stages. A ramjet achieves efficiency by means of high mach to achieve that compression.

In a one-stage design, achieving a 20:1 compression ratio is "easy": just increase the velocity. When we do so, however, we find that achieving that high of a compression ratio requires a mach turbine (airflow over the turbines is mach), which introduces all kinds of wonderful problems.

Grossly inefficient. Very impractical.

But possible.

Again, we do not use multiple stages because achieving high compression with one stage is impossible. We do so because using multiple stages is much more efficient.
 
  • #16
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Absolutely. As I said, it's not impossible. Merely grossly inefficient, and therefore impractical.

A similar, and familiar approach would involve a ramjet. Below Mach 0.5, they have almost zero thrust due to poor compression ratios, and below 600 kts, they're grossly inefficient. Between Mach 2 and 4, however, they'll outperform any turbojet.

The issue is one of compression. In a turbojet, we obtain efficiency by using multiple compression stages. A ramjet achieves efficiency by means of high mach to achieve that compression.

In a one-stage design, achieving a 20:1 compression ratio is "easy": just increase the velocity. When we do so, however, we find that achieving that high of a compression ratio requires a mach turbine (airflow over the turbines is mach), which introduces all kinds of wonderful problems.

Grossly inefficient. Very impractical.

But possible.

Again, we do not use multiple stages because achieving high compression with one stage is impossible. We do so because using multiple stages is much more efficient.
yeah..i agree with you...thanks
 
  • #17
FredGarvin
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Absolutely. As I said, it's not impossible. Merely grossly inefficient, and therefore impractical.

A similar, and familiar approach would involve a ramjet. Below Mach 0.5, they have almost zero thrust due to poor compression ratios, and below 600 kts, they're grossly inefficient. Between Mach 2 and 4, however, they'll outperform any turbojet.

The issue is one of compression. In a turbojet, we obtain efficiency by using multiple compression stages. A ramjet achieves efficiency by means of high mach to achieve that compression.

In a one-stage design, achieving a 20:1 compression ratio is "easy": just increase the velocity. When we do so, however, we find that achieving that high of a compression ratio requires a mach turbine (airflow over the turbines is mach), which introduces all kinds of wonderful problems.

Grossly inefficient. Very impractical.

But possible.

Again, we do not use multiple stages because achieving high compression with one stage is impossible. We do so because using multiple stages is much more efficient.
It is grossly apparent that you have never even seen a compressor in real life. I would suggest you read Hill and Peterson. You have linked to it. Try reading it especially section on axial compressors.
 
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  • #18
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It is grossly apparent that you have never even seen a compressor in real life.
Gee, Fred, I must have missed 'em sitting out there on the wings as I tootled along through the sky. I always wondered what the throttle body was connected to... Thanks for clearing that up!

I'm a retired USAF officer with 2,400+ flight hours and an aero engineering degree from Virginia Tech.

You have yourself a nice day! :biggrin:
 
  • #19
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Gee, Fred, I must have missed 'em sitting out there on the wings as I tootled along through the sky. I always wondered what the throttle body was connected to... Thanks for clearing that up!

I'm a retired USAF officer with 2,400+ flight hours and an aero engineering degree from Virginia Tech.

You have yourself a nice day! :biggrin:
During those 2,400+ hours, were you designing jet engines? That is not to say we don't appreciate your service or your piloting knowledge, but please do not think that means you outclass Fred in jet engine design - it's a bit rude, to say the least.
 
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  • #20
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During those 2,400+ hours, were you designing jet engines? That is not to say we don't appreciate your service or your piloting knowledge, but please do not think that means you outclass Fred in jet engine design - it's a bit rude, to say the least.
https://www.physicsforums.com/showpost.php?p=2509216&postcount=13"I had actually acknowledged Fred's expertise. However, I also pointed out his entering assumption was limited to modern jet engines, designed for their overall efficiency, rather than expanded to consider the possibility (not the practicality) of the OP's question.

Regardless of Fred's credentials, he continues to dodge the issue while resting on laurels. To top it off, https://www.physicsforums.com/showpost.php?p=2512960&postcount=17", was rude and demeaning.

Fred can fend for himself, Cyrus. Whether he chooses to admit his error on the possibility vs practicality issue or not is up to him.
 
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  • #21
minger
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To bring another discipline in here, I'm not so sure a single spool would be possible mechanically. Problems arise as it is seeing pressure ratios in the teen range. That's a lot of force to put on the blade roots...
 
  • #22
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To bring another discipline in here, I'm not so sure a single spool would be possible mechanically. Problems arise as it is seeing pressure ratios in the teen range. That's a lot of force to put on the blade roots...
Again, it's possible, not practical (efficient). Obviously the blades of such an engine would have to be far more sturdy than on typical multi-stage compressors.

As for pressure ratio concerns, compare water, with it's density 1000 kg/m^3 density, and air, with it's sea-level, 20 deg C density of 1.2 kg/m^3 (more than 800 times greater). Yet we use water turbines all the time which handle stresses far exceeding that of a hypothetical single-stage air compressor.

Turbine blade strength isn't an issue, as they can be designed to be as strong as would be required.
 
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  • #23
minger
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From what I understand, reaction water turbines are typically only used in low head appliactions. I doubt they rotate anywhere near the speed that gas turbines do either.

I could fly by flapping my arms if I really wanted to. It would be very practical or efficient, but it could work.
 
  • #24
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From what I understand, reaction water turbines are typically only used in low head appliactions. I doubt they rotate anywhere near the speed that gas turbines do either.
Due to the momentum and viscosity of water, they do not need to.

I could fly by flapping my arms if I really wanted to. It would be very practical or efficient, but it could work.
Well this is a rather inane remark. Since you chose not to address the science/physics/engineering issue directly, I'll respond in kind: Good luck, but please don't overtrust your theory launch yourself off a cliff!

I'm flummoxed by your, Fred's, and others' errant belief that I am in any way proposing single-stage designs, particularly given my oft-repeated "possible, but not practical" mantra.

However, I am equally opposed to the hip-shooting "nope, can't be done" engineering nonsense that's been shoveled into this thread. What part of "possible, but not practical" are you and others failing to understand? Are you errantly believing I'm proposing single-stage designs?

Fess up!
 
  • #25
Astronuc
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I'm flummoxed by your, Fred's, and others' errant belief that I am in any way proposing single-stage designs, particularly given my oft-repeated "possible, but not practical" mantra.

However, I am equally opposed to the hip-shooting "nope, can't be done" engineering nonsense that's been shoveled into this thread. What part of "possible, but not practical" are you and others failing to understand? Are you errantly believing I'm proposing single-stage designs?
From what I read, no one is asserting that one is proposing a single-stage compressure design with a high compression ratio. However, I believe one asserted that it is possible. Please provide an example of such an axial compressor, or failing that, please provide the equation for such a compressor.

Aircraft compressor stage compressor ratios might go to 4 or 5:1, but in the example cited earlier (http://en.wikipedia.org/wiki/Axial_compressor#Spools), the spool engines still have multiple stages, each with a compression ratio ~2:1. The wikipedia article is a bit misleading, because the RB.211 series is a three spool design (first run in Aug 1968). The RB.211-22 has a Single-shaft fan, 7-stage i-p compressor, 6-state h-p compressor, single-stage h-p and i-p, three-stage l-p turbine. Later designs in the series have a 6-stage i-p compressor. The total compression ratio is ~27:1. This is the point that FredGarvin was making.
http://en.wikipedia.org/wiki/Rolls-Royce_RB.211#RB211-22_series_2
http://www.flightglobal.com/pdfarchive/view/1969/1969 - 1205.html
http://www.flightglobal.com/pdfarchive/view/1968/1968 - 0493.html

In a one-stage design, achieving a 20:1 compression ratio is "easy": just increase the velocity.
I'm not so sure it is "easy". Please provide an example or the equations for an axial compressor stage.

There are numerous design constraints such at the blade tip speed and the speed of sound in the compressed air. The back pressure on a single stage with a compression ratio of 20:1 would be enormous and the rotation of the flow would seem to be problematic.
 
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