Some large jet engines have a high combustion temperature to increase maximum thrust. First vanes and blades of the turbine must be cooled to prevent them from melting. One method is to circulate cold air from the compressor inside the element and/or create a coating of air on the element faces using channels and holes.

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Trent turbine HP blade with cooling channels, source

How is air from the compressor channeled to the vanes and blades? How is cooling air prevented to escape to the space between rotor disks instead of going to the cooling channels, under turbine gases pressure?

What approximate fraction of core power is used for such cooling?

  • 1
    $\begingroup$ I don't think you need a high pressure difference. By approximation, the combustion chamber is constant pressure (subsonic flow!) and the faster moving fluid due to combustion creates a lower dynamic pressure (Bernoulli). I'm not sure about the specifics of this system though $\endgroup$
    – Sanchises
    Jan 4 '17 at 19:55

Basically, the 'cooling' air is obtained from the diffuser and/or combustion liner (it will differ in different engines) and is routed both externally and internally (i.e through the casing and spindle) to the nozzle and rotors. Though the details differ from engine to engine, the basics are similar, for example to the NASA Energy Efficient Engine. From NASA CR- 167955 High Pressure Turbine Test Hardware Detailed Design Report:

The Stage 1 nozzle is cooled by air extracted from the inner and outer combustion-liner cavities. The vane leading-edge cavity is fed from the inner flowpath, and the aft cavity is fed from the outer flowpath ...

The Stage 1 rotor is cooled by air extracted at the diffuser mean line.

The image below shows the schematic of the airflow inside the engine used for cooling.

Engine cooling

Engine cooling air supply; image from NASA CR- 167955 High Pressure Turbine Test Hardware Detailed Design Report

Latter stages are cooled with air derived from higher compressor stages. For example, in this engine,

The Stage I shroud is cooled with air supplied from the compressor dis- charge.


The Stage 2 nozzle cooling air is extracted from the compressor at the exit of the seventh-stage stator.

As can be seen in the figure, the cooling air is prevented from escaping back using a number of seals.


Engine seals; image from NASA CR- 167955 High Pressure Turbine Test Hardware Detailed Design Report

  • $\begingroup$ Interesting linked document, thanks. $\endgroup$
    – mins
    Jan 23 '17 at 21:24

To build upon a few points in Peter's answer:

This diagram will probably help visualize it (I copied from this question How is the central hub / shaft casing of a two-spool jet engine assembled?):

GE 90 airflow diagram

The seal between the disks is just a labyrinth seal. Those chambers, however, are pressurized (by similar mechanisms as the cooling air was delivered to the blade). So there is a tiny positive flow of air out of that chambers, which prevents the hot combustion gases from flowing in. Refering to the diagram, the space in front of the 1st stage disk is pressurized from air coming from the stg 10 compressor through the series of passages colored in orange. The space between HPT stage 1 and stage 2 is pressurized by air the comes along the inner diameter of the HP rotor, from an impeller between stage 7 & 8 compressor (in blue), and the space aft of the HPT stage 2 is pressurized from air the comes from the stage 4 compressor (by means of pipes on the outside of the engine, light blue color).

Pretty much every cavity/chamber inside the engine is pressurized somehow, to keep air from going in undesired directions.

I don't know exactly how much air is used for cooling. It's on the order of 1% (i.e. I know it's less than 10%, but certainly more than 0.1%)


Since the pressure of cooling air must be higher than the local pressure in the turbine, this air needs to come from the diffusor area, which is where the highest pressure in a gas turbine is found (see here for details). Normally, it flows through the hollow high pressure spindle so it arrives at the root of the turbine blades from where it finds its way to the surface of the blade. Centrifugal forces help to push the air through the narrow holes, so in effect the turbine blades suck the air in at their root. The seal between the disks is (to my knowledge - I am on shaky ground here) nothing better than a labyrinth seal.

  • $\begingroup$ Between the discs there's isn't any special seal. In some older engines, the surfaces are made flat (typically 50 microns) and bolted together with anything from 50-150 bolts depending on the engine. In newer engines they're welded together. Labyrinth seals are used between objects rotating at different speeds (I.e. Between shafts) $\endgroup$ Jan 5 '17 at 9:01
  • $\begingroup$ @Notts90: Welding will make maintenance impossible, so the whole turbine section must be replaced in one go. Is that really so? $\endgroup$ Jan 5 '17 at 12:11
  • $\begingroup$ I know at least one major manufacturer welds some of the compressor discs to make a set of 2-3 stages in one of the newer engines. I don't think its uncommon to have 2-3 stages permanently joined together. A different manufacturer actually machines 2 stages inc. blades from a solid piece of metal and that's an old engine. $\endgroup$ Jan 5 '17 at 12:18
  • $\begingroup$ @Notts90: Ah, in a compressor this makes more sense. The fully machined stages are called blisks, as in bladed disks. $\endgroup$ Jan 5 '17 at 17:53
  • $\begingroup$ They are indeed! $\endgroup$ Jan 5 '17 at 18:12

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