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I thought that the velocity going through a turbojet increased as it went through the compressor and the burner. Intuitively, if the pressure/enthalpy increases, so would the velocity, right? I came across with this image, though:

Magnitudes-Through-Turbojet

As you can see, this image says that overall velocity actually decreases at the exit of both compressor and burner.

What is the logical explanation behind this? Shouldn't the compressing/burning effect actually accelerate the flow?

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  • $\begingroup$ The law of conservation of energy dictates that the pressure and velocity of a fluid are inversely proportional. So the higher the pressure the lower the velocity and vice versa. $\endgroup$
    – TomMcW
    Jun 13, 2016 at 18:15
  • $\begingroup$ I left temperature out of that, but it still applies. Compressing air reduces its velocity, accelerating air reduces its pressure. $\endgroup$
    – TomMcW
    Jun 13, 2016 at 18:18
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    $\begingroup$ @TomMcW, there is an anomaly in the graph though. Mostly is shows the speed decreasing as the pressure increases and vice versa, but in the combustor stangely both pressure and velocity decrease, while temperature remains constant. $\endgroup$
    – Jan Hudec
    Jun 13, 2016 at 18:33
  • $\begingroup$ @JanHudec That is odd. Neither this graph, nor this one nor this one show that, so I'm wondering if the graph has an error. Dunno tho. $\endgroup$
    – TomMcW
    Jun 13, 2016 at 18:46
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    $\begingroup$ @TomMcW, well, at the exit of compressor it does decrease, because the combustor is wider than the compressor, so the velocity must decrease from conservation of mass. According to the graphs you found, the velocity increases throughout the burner (combustor), so that appears to be an error in the graph. $\endgroup$
    – Jan Hudec
    Jun 13, 2016 at 18:51

2 Answers 2

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The flow enters the compressor at approximately Mach 0.4 to 0.6 in order to ensure a high mass flow with still subsonic compressor blade speeds. Compressing the air allows to reduce flow speed, and between compressor exit and combustion chamber is a diffusor to reduce flow speed even more. Why? To ensure a high degree of combustion! The longer the reactants remain in the combustor, the better. And large combustors are heavy, so they are kept as short as possible.

As soon as the air heats up, it picks up speed. In that respect your graph is not quite correct (or it assumes a rapidly widening cross section through the combustor). But the general relation between pressure and speed is right: While no energy is added or removed, the sum of pressure energy and kinetic energy of the flow is constant. While both the compressor and the combustor add energy and the turbine removes it again, the nozzle is such a region of constant energy. Here the remaining pressure of the flow is converted into speed.

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This image that TomMcW posted in the comments is a much better representation of what is going on.

enter image description here

First, looking at axial velocity only is only half of the story, through the compressor and especially through the turbine the airflow can be more circumferential than axial. In fact the point of the stator vanes and turbine nozzles (turbine vanes) is to turn the airflow to give the proper angle of attack on the blades.

The velocity drops in the compressor because the flow area doesn't decrease as quickly as the air density increases, by design. This increases the efficiency of the compressor and keeps the air moving slower at the compressor exit.

The flow area increases a lot at the compressor discharge going into combustor, this is to slow the air down for better combustion and to ensure the static pressure on the outside of the liner is higher than on the inside. In the combustor, two things happen to increase the velocity, first a massive amount of heat is added, decreasing the density requiring a higher velocity and the area gets smaller, also increasing velocity, both can be calculated with the continuity equation. You also get a small amount of mass increase through the combustor from the fuel.

Now at the exit of the combustor you have the Stage 1 turbine nozzles, which accelerate the airflow to Mach 1, and generally with a very sharp circumferential angle. The air then slows and is straighten as the turbine blade extracts work. This is repeated through each stage of the turbine, although it will not achieve Mach 1 again.

Also remember Mach 1 changes throughout the engine with total temperature.

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  • $\begingroup$ Your answer was equally as good as that of @PeterKampf, thank you both for your insights! $\endgroup$
    – Jose Lopez Garcia
    Jun 14, 2016 at 18:46

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