GB2356671A

GB2356671A - Gas turbine engine cooling

Info

Publication number: GB2356671A
Application number: GB9927957A
Authority: GB
Inventors: Neil Howes
Original assignee: Alstom Power UK Holdings Ltd
Current assignee: Alstom Power UK Holdings Ltd
Priority date: 1999-11-27
Filing date: 1999-11-27
Publication date: 2001-05-30
Anticipated expiration: 2019-11-27
Also published as: GB9927957D0; GB2356671B

Description

2356671 GAS TURBINE ENGINE This invention concerns a gas turbine engine,

and more particularly concerns cooling of one or more vanes of turbine means included in said 5 engine.

European patent application published under No. EP 0 656 468 discloses a gas turbine engine comprising air compressor means to produce compressed air, turbine means to be driven by hot products of combustion and providing rotary drive to drive said compressor means, and combustor means to burn fuel utilising aforesaid compressed air as combustion air in the production of said hot products of combustion. This gas turbine engine further comprises an annular chamber surrounding the axis of the engine, compressed air from a final stage of the air compressor means discharging into the chamber and being mainly used as combustion air by a plurality of fuel burning combustors disposed in the chamber and directing the hot products of combustion through a first stage of stationary vanes of the turbine means. A portion of the compressed air is bled-off from the chamber through a pipe and supplied as cooling air passing through passages in the said stationary vanes, and then returned to the chamber. To ensure a suitable differential pressure to drive the cooling compressed air through the stationary vanes and back to the chamber, the pipe includes a fan to propel the air. Apart from the addition of the extra fan adding complexity, should the fan fail there is a risk that cooling of the aforesaid stationary vanes will be lost because of loss of differential pressure to circulate the cooling air and thus there is a risk of rapid engine failure shortly after fan failure.

According to the invention a gas turbine engine comprises air compressor means to produce compressed air, turbine means to be driven by hot 2 products of combustion and providing rotary drive to drive said compressor means, combustor means to burn fuel utilising aforesaid compressed air as combustion air in the production of said hot products of combustion, said compressor means comprising at least first and second stages wherein the first stage is earlier than the second stage and compressed air produced at the first stage is subsequently compressed to a higher pressure at the second stage, and at least a part of the compressed air output from the second stage being supplied as cooling air to have cooling effect on one or more vanes or blades of said turbine means and being returned to the first stage.

Because the cooling air can be circulated by the pressure difference between the second and first stages of the air compressor means, no additional fan or pump need be provided. Also the peak operating pressure of the said cooling compressed air can be substantially equal to the discharge pressure of the air compressor means, when the said second stage is the final stage in the air compressor means. Thus the chance of increased pressure losses, for a given design of engine, due to an increase in the pressure of the cooling air above that at which it is discharged from the compressor means does not arise.

Preferably said cooling air is returned from said one or more vanes or blades of the turbine means through heat exchange means to the first stage, and heat is extracted from the cooling air at the heat exchange means. By cooling the cooling air prior to reintroducing it to one or more later stages of the air compressor means, an inter-cooler has in effect been introduced into the system and this may improve performance of the air compressor means.

3 Fluid coolant may pass through the heat exchange means to extract heat from the cooling air. The fluid coolant may be or comprise aforesaid fuel and/or may be or comprise an additive, for example water, supplied to the combustor means.

The or at least one of said vanes or blades may be a stationary vane or a stator blade. For example, the or at least one of the vanes may be a stationary nozzle guide vane.

Alternatively or additionally, the or at least one of said blades may be a rotor blade.

The second stage of the compressor means can be a final stage thereof.

The invention will now be further described, by way of example, with reference to Figure 1 of the accompanying drawing which is a diagrammatic and fragmentary representation of a gas turbine engine formed according to the invention.

Referring to Figure 1, a gas turbine engine 2 comprises a turbine section 4 comprising rows of stationary nozzle guide vanes and stator blades alternating with rows of rotor blades (one such rotor blade in a row being indicated at 6a and another such rotor blade in a subsequent row being indicated at 6b) mounted on a rotating shaft 8 having mounted thereon rotating vanes of an air compressor section 10 comprising stationary vanes alternating with the rotating vanes of the compressor section. Compressed air issues from an end 12 of the compressor section 10, at a final compressor stage of the compressor section, into an annular chamber 14 surrounding an axis of the engine and containing a plurality of combustors (one shown at 16) arranged about the axis. Some of the issuing 4 compressed air follows a path such as indicated by arrow 18 and is utilised as combustion air supporting the burning of fuel in the combustors 16 producing hot products of combustion which drive the rotor blades of the turbine section 4 on entering the latter through an aforesaid row of stationary nozzle guide vanes (one such stationary nozzle guide vane being indicated at 20a) forming an inlet to or first stage of the turbine section; an aforesaid stator blade in a row of stator blades being indicated at 20b.

Other of the issuing compressed air from the compressor section 10 follows a path such as indicated by arrow 22 and enters a passage 24 leading to a radially inner annular manifold 26 feeding the compressed air through internal passageways in the stationary nozzle guide vanes 20a of the annular first stage row of stationary vanes of the turbine section so that the air emerges from the stationary vanes 20a at their radially outer ends into a radially outer annular manifold 28. From manifold 28 the air enters a pipe or passage 30 which includes heat exchanger 32. At 34 an end of the passage 30 opens into the air compressor section 10 at an earlier compressor stage where the pressure of the compressed air produced thereat has a lower pressure than that emerging from the final compressor stage at 12. Thus the pressure differential between the compressed air at the later and earlier stages 12 and 34 causes compressed air to follow path 22 to flow through passage 24, stationary vanes 20, and passage 30 and act as cooling air on those vanes to carry heat therefrom and yield it to fluid coolant flowing in path 36 through heat exchanger 32.

The amount of compressed air entering the passage 24 may be any desired amount. For example up to substantially 10% of that issuing from final compressor stage 12. Passage 30 may include airflow direction control means 38, for example valve means, to ensure the flow of cooling air is always in direction 40.

The fluid coolant flowing in path 36 along direction 42 may be liquid or gas. For example, the fluid coolant may be fuel which is heated in the heat exchanger 32 before being supplied on path 44 to the combustors 16 for burning.

Additionally or as an alternative, the fluid coolant may be an additive heated in the heat exchanger 32 and supplied on path 42 to the combustors. For example, the additive may be water which may be used for NOx reduction.

Additionally, or as an alternative to the stationary nozzle guide vanes 20a, one or more stator blades in the turbine section 4 may be supplied with cooling air by passage 24 from which stator blade(s) the air is carried away by passage 30. In a further addition or alternative, the passages 24 and 30 may be arranged to convey cooling air to and carry it from one or more rotor blades or rotating parts of the turbine section 4.

The cooling air may be directed onto inner surfaces of the vanes or blades of the turbine section 4 as jets of air impinging on those inner surfaces (impingement cooling). In addition, bleed passages may lead from an interior of any such vane or blade to open at its outer face so that some of the cooling air bleeds from the interior of the vane or blade to flow as a cooling film over the outer face (film cooling) which means some of the cooling air is lost.

If desired, the cooling compressed air can be taken from a stage of the 30 compressor 10 which is earlier than the final stage at 12, provided the 6 stage from which the cooling compressed air is taken is later than the stage at 34 and is thus at higher pressure than the air at stage 34.

Also, if desired, heat exchanger 32 may be omitted.

Claims

7 CLAIMS

1. A gas turbine engine comprising air compressor means to produce compressed air, turbine means to be driven by hot products of combustion and providing rotary drive to drive said compressor means, combustor means to burn fuel utilising aforesaid compressed air as combustion air in the production of said hot products of combustion, said compressor means comprising at least first and second stages wherein the first stage is earlier than the second stage and compressed air produced at the first stage is subsequently compressed to a higher pressure at the second stage, and at least a part of the compressed air output from the second stage being supplied as cooling air to have cooling effect on one or more vanes or blades of said turbine means and being returned to said first stage.

2. A gas turbine engine as claimed in claim 1, in which said cooling air is returned from said one or more vanes or blades of the turbine means through heat exchange means to the first stage, and heat is extracted from the cooling air at the heat exchange means.

3. A gas turbine engine as claimed in claim 2, in which fluid coolant passes through said heat exchange means to extract heat from said cooling air.

4. A gas turbine engine as claimed in claim 3, in which the fluid coolant is or comprises aforesaid fuel.

5. A gas turbine engine as claimed in claim 3, in which the fluid coolant is or comprises additive supplied to said combustor means.

8

6. A gas turbine engine as claimed in claim 5, in which said additive is water.

7. A gas turbine engine as claimed in any one preceding claim, in which the or at least one of said vanes or blades is a stationary nozzle guide vane.

8. A gas turbine engine as claimed in claim 7, in which the or at least one of said vanes or blades is a stator blade.

9. A gas turbine engine as claimed in any one of claims 1 to 6, in which the or at least one of said vanes or blades is a rotor blade.

10. A gas turbine engine as claimed in any one preceding claim, in which the or at least one of said vanes or blades comprises internal passage means through which aforesaid cooling air flows.

11. A gas turbine engine as claimed in any one preceding claim, in which the second stage is a final stage.

12. A gas turbine engine as claimed in any one preceding claim, in which the second stage discharges compressed air into a chamber including said combustor means, and the cooling air is taken from the chamber.

13. A gas turbine engine substantially as hereinbefore described with reference to the accompanying drawing.