When a jet aircraft flies at maximum speed, are compressor blades actually a barrier to the air flow, or is the airflow accelerated by the compressor?

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    $\begingroup$ Hello Yayang, welcome to aviation.SE! I tried to rewrite your question in such a way that it is easier understood. Please improve my changes if you want to. I have removed your second question; please only ask one question at a time. You can ask it as a separate question. $\endgroup$ – DeltaLima Mar 11 '17 at 14:15
  • $\begingroup$ I think it's wrong to think of the compressor as a barrier, your taking air and forcing it into a high pressure area. $\endgroup$ – Notts90 Mar 11 '17 at 16:29
  • $\begingroup$ Note that @DeltaLima dropped the second question. On this site if you want to ask two question, it is strongly preferred to make two posts. $\endgroup$ – Jan Hudec Mar 12 '17 at 17:57


These are not exhaustive alternatives. Keep in mind that subsonic fluid flow when constricted accelerates and pressure decreases (Bernoulli's principle).

What the compressor does is increase the pressure. The speed decreases slightly.

are compressor blades actually a barrier to the air flow

Actually it's complicated. In a sense they are. The compressor always takes the air at around M0.5 irrespective of flight speed. That means at slower speed the engine needs to suck air from wider stream while at higher speeds some air is spilled over the intake lips. So in a sense, the compressor is a barrier.

But then, the compressor must increase the energy of the stream, otherwise the engine could not produce any thrust. Just the energy goes in pressure, not velocity. So in this sense it is definitely not a barrier.

In fact, the thrust acts mainly on the compressor, because that is the most significant aft-facing surface on which the pressure in the combustion chamber can act.

At supersonic speed it works differently. Now pressure changes can't affect the flow upstream, so the air can't be spilled over the intake lips anymore. Instead now the intake must decelerate it by shock waves and then expansion in a diffusor.

This is needed, because turbomachinery only works well in subsonic flow¹. Since pressure increases as the flow speed decreases, this also increases the overall pressure ratio of the engine and thus its efficiency.

Note that the air also heats up, via adiabatic heating, as it slows down and compresses. Somewhere around M5–6, the air becomes so hot from the compression that adding more heat to it by burning fuel is no longer practical, so above that speed, the combustion must occur in supersonic flow. This is the scramjet engine. It does not have any compressor or turbine, because it gets all the compression from constricting the flow and it turbomachinery wouldn't work at that speed anyway.

  • $\begingroup$ Should the compressor be considered a barrier, or is it instead the intake system that slows the initial ram airflow? $\endgroup$ – J Walters Mar 13 '17 at 10:34
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    $\begingroup$ @JonathanWalters, at subsonic speed, the compressor, but really the whole engine, is what causes the pressure increase ahead of it. At supersonic speeds pressure changes don't propagate upstream, so there it is the intake shape (supersonic intakes are low while subsonic ones are usually very short) that slows and precompresses the flow. $\endgroup$ – Jan Hudec Mar 13 '17 at 17:50

Jet engine axial type compressors rotor blades do accelerate airflow. It is important to note that the net speed change in velocity is usually negative, that is, airflow through the compressor section taken as a whole is usually slowed. Additionally, the engine intake generally slows airflow as well while increasing pressure. This occurs within the airflow prior to direct contact with the compressor.

Jet engine axial compressors are comprised of multiple pressure stages with the primary design of increasing pressure, not velocity. Each pressure stage is comprised of pairs of rotors and stators. The rotors consist of rows of compressor blades arranged radially on a rotating spindle. The stators consist of fixed rows of vanes arranged radially and placed after and between the rotors.

The compressor blades of the rotors accelerate the airflow and impart a momentary velocity increase. Subsequently, the stator vanes and the decreasing geometry of the compressor design slow the speed and change the airflow direction (which is also an acceleration). Each pressure stage raises the pressure, typically with a small net decrease in speed.

See the following figure and text excerpt from the Jeppesen A&P Technician Powerplant Handbook:

The task of an axial compressor is to raise air pressure rather than air velocity. Therefore, each compressor stage raises the pressure of the incoming air while the air's velocity is alternately increased then decreased as airflow proceeds through the compressor. The rotor blades slightly accelerate the airflow, then the stator vanes diffuse the air, slowing it and increasing the pressure. The overall result is increased air pressure and relatively constant air velocity from compressor inlet to outlet.

enter image description here Figure 3-29

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    $\begingroup$ I would be interested in understanding why this was downvoted. Have I missunderstood the question? $\endgroup$ – J Walters Mar 11 '17 at 20:54
  • $\begingroup$ I am not the downvoter. Just a remark, that R/S velocity plot looks strange? Shouldn't be one accelerating and the other decelerating (with corresponding pressure variations) those discontinuities look a bit odd to me. see this answer (but is is about turbines, you have to "revert it") aviation.stackexchange.com/questions/20785/… $\endgroup$ – Federico Mar 12 '17 at 20:45
  • $\begingroup$ I was trying to explain the rotor/stator concept to my vehicle mechanic at work to explain why a standard ventilation fan is going to be good at accelerating air but lousy at creating pressure because it lacks stators.. I wasn't doing so well with my explanation and he wasn't buying it. Maybe I can show him your answer. $\endgroup$ – TomMcW Mar 13 '17 at 3:30
  • $\begingroup$ @Federico Yes, that's true; I hadn't noticed that. The labeling should instead indicate individual pressure stages. Not my work, straight out of the textbook. Perhaps I should fix it. $\endgroup$ – J Walters Mar 13 '17 at 3:53
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    $\begingroup$ Following @Federico 's remark, here is a drawing from RR, for the compressor section. Source. $\endgroup$ – mins Mar 13 '17 at 10:29

Well some basic math might reveal the answer.

If a turbojet has a compression ratio of say 20:1, gas flow entering the combustion chamber will be at a 20 times higher pressure than the gas flow entering the engine inlet.

If we assume the overall cross sectional area of the gas passage through the stators and rotors in the compressor does not change (this is usually not the case) then the gas flow would have to slow.

Since Mdot = rho * Vdot = rho * A * vdot

and by IGL for simplicity, P = rho * R * T

this gives

Mdot = (P/RT) * A * vdot

Since Mdot and A remain constant and P increases, vdot must decrease to compensate. While temperature does increase, it is not enough to compensate for the change in pressure, which can be shown with a sensitivity analysis.

Ideally you want to decelerate the gas flow as little as possible as the decrease in momentum creates a drag impulse upon the engine. Both centrifugal and axial ompressors narrow in volume as gas flows through the core to reduce this.

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    $\begingroup$ I don't know of any turbine engines with axial flow compressors with a gas passage that does not decrease in the direction of flow. Usually the decrease is quite substantial. $\endgroup$ – J Walters Mar 11 '17 at 20:07

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