Pressure Ratios

KJ_Lesnick

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As a general rule, a high pressure ratio means more thrust and more efficiency: I'm curious if there's a pressure ratio that would reach a sufficient level that efficiency would actually fall off


It sounds bizarre, but I do vaguely remember a primitive engine simulator NASA had on its site and when I set PR's above 50 performance fell off rather than go up
 
Next-gen civil engines are aiming for ratios > 60, or even 70, but I think the main constraint is the turbine entry temperatures that result,
 
How is pressure ratio determined? I'd always thought it was the amount of compression the engine provides between entry and the 1st stage of the turbine. But then I've also heard it's the difference in pressure between the air entering the engine and exiting the nozzle. ???
 
For the engine's thermodynamic cycle, i think what matters is the ratio of P3 (usually the total pressure at the exit of the last compressor stage) to P0 (freestream total pressure). This is in turn the product of the total pressure ratios between freestream/inlet throat, inlet throat/fan face, and first (and optionally second) compressor stages.


There is a little bit of drop in the burner, but that's essentially the raise in pressure being fed to the turbine, where most of the pressure rise is brought back down depending on engine application (turboprop, turbojet, etc...).
 
Sferrin

1. When the term pressure ratio is said, it usually means the pressure at the last compressor stage relative to the airflow in front of the engine if I recall

2. EPR means the pressure of the air entering and exiting the engine


Kiltonge

Next-gen civil engines are aiming for ratios > 60, or even 70, but I think the main constraint is the turbine entry temperatures that result
I didn't know they were aiming that high.

I was under the impression that if you went high enough you'd end up with some of the various problems (could be wrong)
  • As you compress air, just like compressing a spring: It becomes increasingly harder to "squeeze" it together without changing the density of the flow (i.e. cooling it down)
  • More energy is required to squeeze the air together harder; ram-compression results in drag increase that increase out of proportion to thrust increases)
 
Pressure ratio is just the difference between the front and the back of something. It can be specified for each compressor stage, or for the whole compressor. For the whole engine end-to-end, thermodynamic efficiency requires the pressure ratio - and hence the exit speed - should be as low as possible. The most meaningful expression for a whole engine is probably that between ambient and the combustion chambers. Even a turbine-free ramjet has such a pressure ratio.

As the ratio increases, you get more oxygen in there to burn more fuel and drive harder backwards against the turbines, and engine efficiency improves. But if you go over the top, the pressure pushes back too hard against the incoming airflow and the thing stalls or chokes.
 
steelpillow

1. So the excessive pressure poses a problem more from a stall-standpoint than net-thrust?

2. What factors allow the highest pressure without stall?
 
1. Yes, internal overpressure causes the intake to stall and air flows round the duct instead of into it.

2. Essentially, the sustainable pressure ratio is a function of the compressor stage, i.e. its design and its speed of rotation. Blade tip speed must remain subsonic, which ties the rotation speed to the diameter. Therefore, the design - how many stages and how effective each stage is - becomes the dominant factor. Ram air pressure can help a little, and that is a function of aircraft speed and intake design.
 
steelpillow said:
1. Yes, internal overpressure causes the intake to stall and air flows round the duct instead of into it.

2. Essentially, the sustainable pressure ratio is a function of the compressor stage, i.e. its design and its speed of rotation. Blade tip speed must remain subsonic, which ties the rotation speed to the diameter. Therefore, the design - how many stages and how effective each stage is - becomes the dominant factor. Ram air pressure can help a little, and that is a function of aircraft speed and intake design.

For #2, compressor stall can be sidestepped somewhat if not using bladed airfoils that can stall. The most obvious example being RamGen's rampressor, which is a rotary ramjet throat, which uses ram compression like a ramjet inlet to compress air. Thus the rotary component can exceed the speed of sound (though materials stress/tension limits still limit diameter). Though I suppose the rampressor's inlet lip could have a stall of sorts.
 
Yes, I should have said that I was considering conventional jet engines. Once you go fast enough, the intake geometry and the sheer speed begin to dominate. Faster still and the compression zone outflow becomes supersonic, as in the scramjet.

Rotary airflow is conventionally used to form a centrifugal dust filter in desert conditions and similar, but that is driven by engine suction rather than by ram air.

Bu why would RamGen want to speed up the airflow by spinning it? High flow speeds are usually associated with low compression, and the usual problem is to slow it down enough to compress it enough for an efficient thermodynamic cycle.
 
steelpillow said:
1. Yes, internal overpressure causes the intake to stall
And I assume a partial variable would be the amount of pressure gain per stage?

2. Essentially, the sustainable pressure ratio is a function of the compressor stage, i.e. its design and its speed of rotation.
So each blade has a different limit, or the pressure gain per stage as a general rule plays a role in stalling?

Blade tip speed must remain subsonic, which ties the rotation speed to the diameter.
Actually that's not exactly true: Many aircraft engines have supersonic tip-speeds as the RPM nears and reaches either maximum continuous, military, or climb-power.

The inflowing air is subsonic, at first controlled only by compressor RPM, inlet shape, and wind (if applicable); as you go faster it generally remains subsonic. The combined flow with rotational velocity combines to produce a supersonic flow across the whole blade (presumably this would take hold around 0.30 to 0.70 mach; I don't know if it covers every stage or the first few). If I recall, the compressor blades might have to be properly designed for this to work right, but evidently it increases the maximum pressure-ratio per stage and seems to make the blades more stall resistant (This does *does* surprise me, as critical alphas are generally lower at transsonic speeds than subsonic speeds).

As I understand it, as tip velocity reaches around 1.4 mach, there may be a tendency for performance to fall off.

I should have said that I was considering conventional jet engines.
I was largely talking about turbojet/turbofan-type engines -- this is the subject of interest for now

Once you go fast enough, the intake geometry and the sheer speed begin to dominate.
In terms of compression, absolutely. Ramjets are more efficient evidently it would appear in terms of power-to-weight at least, if not in terms of the amount of pressure produced by decelerating the airflow (on a jet you have a turbine to deal with, on a ramjet -- no need).

Faster still and the compression zone outflow becomes supersonic, as in the scramjet.
As I understand it with this kind of propulsion system the argument is that by decelerating the airflow large amounts results in increased temperatures and pressures.

The pressures require a potentially stronger duct, which weighs more, and thus requires a heavier airframe to bear it; exotic materials, active cooling, and any of the following (though I'm not sure how severe this is) in terms of any of the following: Chemical interaction between O2 and N2; disassociation of O2 to mono-atomic Oxygen; possibly ionization to some degree.

The argument is that it's lighter and more efficient to simply slow the flow from high supersonic/hypersonic to low supersonic speeds as there is still plenty of pressure to be gained this way: It also doesn't require the inlet duct to be convergent-divergent as the flow is never subsonic; the nozzle doesn't require it either as the flow is always supersonic.

The issue usually involves effectively injecting the fuel into the airflow, the traditional spray-bars and flame holders used in ramjets are based on the flow being subsonic. When hit with supersonic flow all sorts of complex shock-wave patterns form. You instead inject off the walls and you're good.


This is largely academic as there have been ways to produce similar or superior efficiencies to ramjets by simply spraying the fuel into the inlet duct substantially upstream of the combustion-chamber: It not only mixes better, possibly vaporizes, but it also absorbs a great deal of heat which is an active cooling effect. Seems simpler to implement.
 
When talking about engine pressure ratios in an engine cycle context they are referring to the ratio of total pressure at compressor delivery to the combustion plenum compared to free stream total pressure. EPR, or "Engine Pressure Ratio" has a decent correlation to engine thrust and is used by some engines and refers to the ratio of Jetpipe total pressure to inlet total. Most modern engines have eliminated EPR because as bypass ratios have risen N1 (LP spool / Fan speed) is also closely correlated to thrust and as you need to measure this anyway you can eliminate the extra parameter. EPR is notoriously famous for causing the Air Florida crash when the inlet total probe iced and caused high EPR readings that caused the pilots to takeoff with the thrust too low.

The reason SFC drops off above a certain limit is because ideally you would like to fully expand the air back to ambient, thus extracting all of the pressure energy from the flow; this is never really fully possible but it becomes very inefficient when too much expansion takes place across the final nozzle.
 
iancheyne said:
When talking about engine pressure ratios in an engine cycle context they are referring to the ratio of total pressure at compressor delivery to the combustion plenum compared to free stream total pressure. EPR, or "Engine Pressure Ratio" has a decent correlation to engine thrust and is used by some engines and refers to the ratio of Jetpipe total pressure to inlet total.
I know vaguely what EPR is... I was talking more about max PR.

Most modern engines have eliminated EPR because as bypass ratios have risen N1 (LP spool / Fan speed) is also closely correlated to thrust and as you need to measure this anyway you can eliminate the extra parameter.
With high-bypass fans...

EPR is notoriously famous for causing the Air Florida crash when the inlet total probe iced and caused high EPR readings that caused the pilots to takeoff with the thrust too low.
Because of the following

1. They did a power-back when they weren't supposed to
2. They turned the engine anti-ice off
3. They didn't return for another de-icing and used the exhaust of a DC-9 in front of them to melt off and blow away the snow on their wings (it didn't work -- it liquified the ice and pushed it back -- it then refroze)

The reason SFC drops off above a certain limit is because ideally you would like to fully expand the air back to ambient, thus extracting all of the pressure energy from the flow; this is never really fully possible but it becomes very inefficient when too much expansion takes place across the final nozzle.
So one limit becomes stall and surge margins, the other is expansion ratios become impractical?
 
KJ_Lesnick said:
As a general rule, a high pressure ratio means more thrust and more efficiency: I'm curious if there's a pressure ratio that would reach a sufficient level that efficiency would actually fall off


It sounds bizarre, but I do vaguely remember a primitive engine simulator NASA had on its site and when I set PR's above 50 performance fell off rather than go up

Sorry for the rather delayed response.

There are a couple of limits on pressure ratio. If you run through a Brayton cycle analysis with realistic efficiencies for the turbomachinery and realistic pressure drops and efficiency for the combustor, you will be able to see them. The optimum pressure ratio for efficiency (which isn't the same as that for optimum pressure ratio for thrust per unit mass of air flow) is dependent on turbine entry temperature. You can find some programs that do this sort of analysis at Jack Mattingly's website, http://www.jsmatt.com/custom1.html
 
Kiltonge said:
Next-gen civil engines are aiming for ratios > 60, or even 70, but I think the main constraint is the turbine entry temperatures that result,
For reference, attached is a link to an Aviation Week article on Rolls Royce's Advance and UltraFan developmental programs for future commercial jet engines. Advance is aiming for an Overall Pressure Ratio (OPR) of 60:1, and UltraFan would be aiming at 70:1.

http://m.aviationweek.com/commercial-aviation/rolls-royce-details-advance-and-ultrafan-test-plan
 
KJ_Lesnick said:
As a general rule, a high pressure ratio means more thrust and more efficiency: I'm curious if there's a pressure ratio that would reach a sufficient level that efficiency would actually fall off


It sounds bizarre, but I do vaguely remember a primitive engine simulator NASA had on its site and when I set PR's above 50 performance fell off rather than go up

There are some practical issues. Obvious ones include weight and cost (note that the J-57 was heavier relative to thrust than other designs that employed lower pressure ratios; an engine with a pressure ration of 50 would be extremely heavy because of the large number of compressor stages and the need for extremely high structural strength to handle these pressures. More subtle issues concern losses from air leaking back across the tips of the compressor blades and the high temperatures which would result. It would seem that there are diminishing returns but good engineering can
 

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