Dismiss Notice
Join Physics Forums Today!
The friendliest, high quality science and math community on the planet! Everyone who loves science is here!

A Jet Engine doubt.

  1. Jun 11, 2007 #1
    A Jet Engine doubt.........

    Hi folks....
    My aerospace propulsion professor taught me that, in the combustion chamber of a jet engine, during combustion, the pressure instead of increasing, decreases.
    I have 2 questions.
    1) Is it true?
    2) If yes, how?

    How and in what way does temperature, and pressure affect the efficiency of a turbine in the jet engine?

    I am counting on you little Einsteins out there........... :smile:
    Last edited: Jun 11, 2007
  2. jcsd
  3. Jun 11, 2007 #2


    User Avatar

    Staff: Mentor

    Let me answer the question with another question: what direction does a fluid flow: from low pressure to high pressure or from high pressure to low pressure?

    As far as efficiency goes, it is a function of pressure: http://www.adl.gatech.edu/classes/propulsion/prop5.html [Broken]
    Last edited by a moderator: May 2, 2017
  4. Jun 11, 2007 #3
    a fluid flows from high pressure to low pressure........
    so what? please tell me...........
  5. Jun 11, 2007 #4
    and u say efficiency of the turbine plainly depends on the pressure alone?
  6. Jun 11, 2007 #5
    Remember pressure is two or tree things, static and dynamic pressure and the sum of dynamic and static pressure.

    I would belive that the static pressure decreases slightly, while the dynamic pressure and the temperature increases, so that the total amount of energy in the exaust gasses increases even though there might be a slightly drop in static pressure. (If I remember it right.)

    The increase in temperature and dynamic pressure, and with other word the increase of energy content is due to the burning fuel.
  7. Jun 11, 2007 #6
    By the way .. I allmost wrong, I forgot one thing.

    Most jet engines use to have an difuser in the area between the comressor outlet and the front end of the combustion chamber. In this area the static pressure goes up and the dynamic pressure goes down.

    Then it is a rather stable area trough the combustion chamber. In the last part of the comustion chamber, the combustion gasses increases velocity, and static presure and temperature goes down, before stream of combustion gasses enters the turbine blades.

    I found this link that tells something about how static pressure and velocity (dynamic pressure) changes trough the jet engine. The picture is not very precise as it does not mention "the spesial things" that happend when the air stream enters and leaves the comustion chamber area. In this way the drawing is a bit "oversimplified".

    http://science.radioelectronics.biz/aeronautics390/jetEngine.htm [Broken]
    Last edited by a moderator: May 2, 2017
  8. Jun 11, 2007 #7


    User Avatar

    Staff: Mentor

    Ok, so for air to flow from the compressor to the combustion chamber to the diffuser, which must be at the highest and which must be at the lowest pressure?
  9. Jun 11, 2007 #8


    User Avatar

    Staff: Mentor

    For the purpose of the OP, we need only discuss the total pressure. By Bernoulli's principle, the total pressure along a streamline is constant. When losses are taken into account, the total pressure will drop as the fluid travels through the duct.
  10. Jun 11, 2007 #9


    User Avatar

    Staff: Mentor

    Did you read what the link said and see how the equations work?
  11. Jun 11, 2007 #10


    User Avatar
    Science Advisor

    In a well designed burner, the heat addition step, in practice, results in a pretty constant pressure exiting the burner and going into the HP turbine. The pressure does drop for a couple of reasons:

    1) It is an unrestrained expansion.
    2) There are flow losses within the burner
    3) There is a decent dilution amount of cooling done to the flow to maintain a cool enough gas temperature to not have too high of an TIT.
  12. Jun 11, 2007 #11
    No, I don't think it will work like that.

    The gas stream trough a combustion chamber does not follow the Bernoulli's equitation at all. This prinsipple will only be valid when the airstream has one fixed amount of energy. In the case of the combustion chamber you add on a lot of burning fuel so that the total pressure will certainly not be constant or decrese.

    It is this increase in total pressure that makes the jet engine to work as a jet engine. If no build up in total pressure, there would in the end (in the exaust nozzle) be no jet trust.

    To make a jet trust the air that passes trough the jet engine will need to have added the dynamic and the total pressure. A jet engine that does not accerlerate the air masses, that will increase the total pressure, will be a shut down jet engine.

    I think I have it approx right above o:)
    Last edited: Jun 11, 2007
  13. Jun 12, 2007 #12

    Remember that the combustion chamber consist of "two halves":

    1. The area outside the internal cobustion liner, that will work allmost as you mention. (Exept for that you have not mentioned the diffusor zone between the compressor area and the combustion chamber, that will increase the static pressure quite a bit.)

    2. Then there is the area (volume) inside the internal combustion liner. Here the static pressure will be allmost a constant or drop slightly, while the dynamic pressure and speed is increasing due to the added amount of energy from the fuel. (And this increase in energy is what makes the engine, in the end, capable of producing trust and work.)
  14. Jun 12, 2007 #13


    User Avatar

    Staff: Mentor

    Have a look at the P-V diagram for a jet engine, Langbein: http://en.wikipedia.org/wiki/Brayton_cycle

    Pressure rises only in the compression stage. Combustion happens at constant pressure (in an idealized engine).

    All thermodynamic cycles are pretty much variations on the same themes. In this case, it is just like how in a steam cycle, water goes into the boiler and is heated at constant pressure.
    Last edited: Jun 12, 2007
  15. Jun 12, 2007 #14
    The p-v diagram of a jet engine menetioned in wikipedia would not be that one you would use if you got an F-16 engine up in the test bed to make a test run. The p-v diagram of wikipedia leaves out a few details, I think.

    One detail that is left out is that the front diffusor of the combustion chamber (the area/voulme between the end og the axial compressor and the front of the combustion chamber inner liner) will work as and should be considered to be a part of the compressor. (So that the compressor actually will consist of one part that is build up with a moving rotor and one part that is buildt up by a static diffusor that continues into the combustion chamber up to the front of the combustion chamber inner liner.

    When you relate to the brayton cycle, these two parts, the dynamic compressor and the static diffusor will be "the compressor".

    (Step 1-2 of the brayton cycle extend into the combustion chamber.)

    The combustion at the approx constant static pressure will take place inside the combustion chamber internel liner.

    (Step 2-3 take place inside the combustion chamber internal liner.)

    As one can se from the bryton cycle at step 2-3 the static pressure is approx constant (innside the inner liner) while the dynamic pressure (the speed) is increasing. As total pressure is the sum of static pressure and dynamic pressure, the total pressure gores up.

    One other thing that is left out from the simplified p-v diagram in wikipedia, is that most real jet engines has a nozzle at the rear og the combustion chamber inner liner and in front of the first stage turbine rotor. This sets up the speed, it reduces the static pressure and increases the dynamic pressure. This will also cool down the air stream so that the first stage turbine is not burned.

    A lot of these smaller details is not possible to read from such a overview that is mentioned in wikipedia.
  16. Jun 12, 2007 #15


    User Avatar
    Science Advisor

    Who mentioned anything about the diffuser?

    The total pressure across a well designed burner is what remains almost constant. If you look at the pressure ratio across a burner, it should be very close to 1.

    Attached is a plot of expected parameters as varying with engine station.
    Last edited: Jun 28, 2007
  17. Jun 12, 2007 #16


    User Avatar
    Science Advisor

    You are absolutely wrong. That is the exact cycle to use as a reference (not that I have ever had a Brayton cycle next to me when we have done an engine test, but who's asking?).
  18. Jun 12, 2007 #17
    Well, I think I'm right. What kind of engine(s) do you test on ?

    What was the test conditions ?

    How do you check what happens inside the combustion chamber ?

    Did you dissassembly the engine to take a look at the different details ?

    Did you then reassemble it, testrun it to send it out do do the flying with pilot and or pax after ?

    Which engine was it ?
    Last edited: Jun 12, 2007
  19. Jun 12, 2007 #18


    User Avatar
    Science Advisor

    If you take a look at my occupation, you will see what I do for a living. I test full up engines as well as engine components for a living.

    My company makes a few different engines for both military and civilian applications.
  20. Jun 12, 2007 #19
    OK, I did, do the same before, but not any more. (Does some other technology stuff now.)

    Used to do F-100/F16 O/H and test drive in teststand and installed on aircraft before. (Worked as a F-16 test engineer) Later on civilian with helicopter and aircraft turbine engines.

    But back to the interesting stuff .. :-)

    Will you really say:

    1. That if you open a jet engine, you will not find a diffusor behind the axial compressor and in front of the combustion chamber inner liner ? And this diffusor will not increase the static pressure into the inner liner ?

    2. In the rear part of the inner liner and in that section that goes over to the turbine area, the air/exaust channel is not formed like a nozzle that increases the exaust gas velocity before entering the turbine area ??!

    I think actually all jet engines I have looked into have been designed like this.

    Which one (engine type/model) is not designed like this, som practical examples ?
  21. Jun 12, 2007 #20
    By the way, I will try to se if I can get up the picture you sendt before.


    If you see the area after the compressor, you can see it is formed as a nozzele. The pressure curve is not accurate because it should show a pressure build up in this area.

    On the other side the increase in speed does near the end of the combustion chamber does show. The drawing is allmiost correct on that one.

    I think this drawing is a good one but it is still a bit to simplified to show all of the prosess. The main thing that is missing is the pressure build up after the last compressor stage.

    By the way, I do agree tha general course books on jet engine theory does explain the pressure diagram for a jet engine like the drawing on your link. But I also think that if you go a bit deeper down on each jet engine model, this courve will look a little bit different.

    .. Just my point of view :-)
    Last edited: Jun 12, 2007
Share this great discussion with others via Reddit, Google+, Twitter, or Facebook