GB2246174A

GB2246174A - Cooled aerofoil for a gas turbine engine

Info

Publication number: GB2246174A
Application number: GB8218492A
Authority: GB
Inventors: Barry William Lings; Susan Marie Allen; Charles Robert Gill; David Anthony Richardson; David Hartshorne
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1982-06-29
Filing date: 1982-06-29
Publication date: 1992-01-22
Anticipated expiration: 2002-06-29
Also published as: GB2246174B

Abstract

A cooled aerofoil, such as the aerofoil of a nozzle guide vane, for a gas turbine engine has a cooling system in which the release of cooling fluid on the convex suction flank is avoided. The cooling system utilises a thin skin 28 which overlays the contoured surface 27 of the convex flank to form a plurality of passageways. A first flow of cooling fluid enters these passages via a plurality of impingement holes 30 and leaves via cross flow passages 31 and 32. The cross flow passages allow the fluid to pass to the opposed, pressure flank of the aerofoil where it leaves the aerofoil via film cooling holes 37. <IMAGE>

Description

A COOLED AEROFOIL FOR A GAS TURBINE ENGINE This invention relates to a cooled aerofoil for a gas turbine engine. In the embodiment illustrated the aerofoil is that of a nozzle guide vane, but the invention could also be applied to the aerofoils of rotor blades and in particular to turbine rotor blades.

There is a continuous search for improved ways of cooling such aerofoils, and as the state of the art advances it often happens that cooling methods are found to be unsuitable for use in various parts of an aerofoil. This has happened with film cooling, which is generally accepted to be a highly efficient form of cooling and which would at first sight be applicable to the whole outer surface of an aerofoil. Unfortunately, the use of film cooling builds up a thicker boundary layer, and this introduces aerodynamic penalties particularly when used on the convex, suction flank of the aerofoil.

Therefore, although it is still acceptable to use film cooling on other parts of the aerofoil surface, it is desirable to use a cooling system for the convex, suction flank which does not involve film cooling or any other cooling system involving the release of cooling fluid into the boundary layer. Clearly, it is then necessary to provide some means for disposing of the cooling fluid used to cool this flank.

The present invention provides a cooled aerofoil in which these problems are solved in an efficient manner.

According to the present invention a cooled aerofoil for a gas turbine engine comprises a thin skin overlying the contoured surface of the convex, suction flank of the vane so as to form therewith a plurality of passages, apertures through which a first flow of cooling fluid may flow into said plurality of passages, at least one crossflow passage through which the first flow of cooling air may flow from said plurality of passages across the section of the aerofoil to the opposed convex pressure flank of the aerofoil, and film cooling holes in said convex pressure flank of the aerofoil and in communication with said cross flow passage so that said first flow of cooling air may leave the aerofoil through said film cooling holes.

Said apertures may be set out to provide impingement cooling of said thin skin.

Said film cooling holes may be formed in a further thin skin which forms at least part of the concave pressure surface of the aerofoil.

There may be a second flow of cooling fluid which flows round said cross-flow passages to the trailing edge of the aerofoil to provide cooling thereof.

The invention will now be particularly described, merely by way of example, with reference to the accompanying drawings in which: Fig. 1 is a partly broken-away view of a gas turbine engine having nozzle guide vane aerofoils in accordance with the invention, Fig. 2 is an enlarged cross-section of one of the nozzle guide vane aerofoils of fig. 1 and in accordance with the invention, Fig. 3 is a further enlarged portion of the section of fig. 2, and Fig. 4 is a view on the arrow 4 of fig. 3 but with the overlying skin removed to expose the contoured surface beneath.

In figure 1 the gas turbine engine comprises a fan 10, an intermediate pressure compressor 11, a high pressure compressor 12, a combustion chamber 13, a high pressure turbine 14, an intermediate pressure turbine 15, and a low pressure turbine 16 all in flow series. Operation of the engine is conventional in that air is taken into the fan 10 where it is compressed. This compressed air is divided into two flows one of which passes between the fan casing 17 and the core engine casing 18 to provide propulsive thrust while the remainder enters the intermediate pressure compressor.

The compressed air is mixed with fuel and burnt in the compressed chamber 13 and the hot gases resulting from this combustion pass through the turbines 14, 15 and 16 to drive them. The turbines are drivingly interconnected with their respective compressors so that they cause them to operate.

Hot gas leaving the turbine 16 provides additional propulsive thrust.

The hot gases produced in the combustion chamber 13 must be directed on to the rotor of the turbine 14 in the correct direction and to this end an annular array of nozzle glide vanes 19 are provided. Each vane 19 comprises an aerofoil 20 and inner and outer platform members. Because the vanes 19 operate in the hottest environment of the engine it is necessary to provide them with some form of cooling and in figure 2 the aerofoil 20 is shown in enlarged sections to illustrate this cooling.

The aerofoil 20 is basically hollow and is divided into forward and rearward cavities by a transverse partition 21.

The forward cavity 22 comprises a simple plenum chamber into which cooling air flows, and which has a plurality of film cooling holes 23 through which the cooling air may flow to cool the leading edge region of the aerofoil.

The rearward cavity 24 is again provided with a flow of cooling fluid, but in this case the cooling system is more complex. The walls of the aerofoil which form the cavity comprise a convex, suction wall 25 and a concave, pressure wall 26. These walls form part of the integrally cast structure of the vane but their outer surfaces do not form the aerodynamic profile of the vane in this region.

Thus the wall 25 is provided with a contoured surface 27 (see also figs. 3 and 4) over which lies a thin, impervious skin 28. The contour of the surface 27 is such as to form a plurality of branched channels 29, which are closed by the skin 28 to produce a branched series of passages. Various of the channels 29 are provided with holes 30 extending through the wall 25 and into communication with the interior of the cavity 24.

The branched passages terminate in two sets of crossflow passages 31 and 32 formed insthe wall 25 which is provided with local projections 33 and 34 to join with the opposed wd?l 26 and within which the cross-flow passages 31 and 32 are formed. The passages 31 and 32 communicate with a further series of channels 35 formed in the surface of the wall 36. As in the case of the channels 29, the channels 35 are closed by a thin skin 36 which overlays the channels and forms the aerodynamic profile of the aerofoil in the region. However, unlike the sheet 28 the sheet 36 is provided ith pluralities of film cooling holes 37 through which cooling air may flow to film cool this surface of the aerofoil.

Operation of the arrangement described so far utilises a first flow of cooling air from the cavity 24. (A flow of air to the cavity 24 is supplied by means not shown but comprising a proportion of the air compressed by the engine which is fed through one or both ends of the aerofoil 20).

The first flow of cooling air passes through the holes 30 to impinge upon the inner surface of the skin 28 and thus to provide some cooling. The air then flows through the branched passages, providing additional convective cooling of the skin 68 and the wall 25. Upon reaching the extremities of te branched passages the air enters the cross-flow passages 51 and 32 and flows across the aerofoil section into the channels 35.

The air provides some additional convection cooling by flown through the channels 35, but its main cooling effect is produced when it leaves the aerofoil via the film cooling holes 37 to filr cool this region of the aerofoil. It will be noted that this film cooling is on the relatively insensitive pressure flank of the aerofoil.

This system through which the first flow of cooling air flows will be seen to satisfy the requirement set out above. Thus it cools the convex, suction surface without the need for films in this critical area and the air used for cooling this area is then re-used to cool the concave pressure surface. This double use of the air uses the air in an efficient manner while solving the problem of dealing with the spent cooling air in a way which does not interfere to any substantial degree with the aerodynamics of the vane aerofoil.

In order to cool the remaining parts of the trailing region of the aerofoil, a second flow of air takes place, from the cavity 24 and in between the cross-flow passages 31 and 32 toward the trailing edge of the aerofoil. This second flow then passel through a further set of impingement holes 3B to impinge upon the undersurface of a rearward extension of the skin 28. A series of chordwise channels 39 are overlain by the skin 28 to form channels, and the second flow of air passes down these channels to exit at the trailing edge 40.

It will be seen that the skin 36 forms part of the boundary of the secondary flow of air in the trailing edge region of the aerofoil. It is clearly possible, if it is found desirable, to provide film cooling holes in this part of the skin 36 to film cool the trailing edge part of the concave pressure surface of the aerofoil.

As described above the invention is illustrated in its application to a noz,le guide vane aerofoil having a cast main portion and overlaying skins to form certain parts of the trailing edge region. However, the invention could clearly be applied to a rotor blade aerofoil, and various methods of manufacture could be used for instance as an integral casting operation to produce the whole aerofoil including the skins and channels.

It should also be noted that the overall thermal performance of the aerofoil may well be improved by the application of a thermal barrier coating made up of a ceramic coat deposited upon a bond coat on the skins and/or remainder of the aerofoil.

Claims

Claims:

1, A cooled aerofoil for a gas turbine engine comprising a thin skin overlying the contoured surface of the convex, suction flank of the vane so as to form therewith a plurality of passages, apertures through which a first flow of cooling fluid may flow into said plurality of passages, at least one cross-flow passage through which the first flow of cooling air may flow from said plurality of passages across the section of the aerofoil to the opposed convex pressure flank of the aerofoil, and film cooling holes in said convex pressure flank of the aerofoil and in communication with said cross-flow passage so that said first flow of cooling air may leave the aerofoil through said film cooling holes.

2. A cooled aerofoil as claimed in claim 1 and in which said apertures are dimensioned and spaced so as to provide impingement cooling of said skin.

3. A cooled aerofoil as claimed in claim1 or claim 2 and in which said film cooling holes are formed in a further thin skin which forms at least part of the concave, pressure surface of the aerofoil.

4. A cooled aerofoil as claimed in claim 3 and in which said further thin skin overlays the contoured surface of the convex pressure flank of the vane so as to form a second plurality of passages through which said first flow of cooling fluid flows before passing through said film cooling holes.

5. A cooled aerofoil as claimed in any one of the preceding claims and in which there is a second flow of cooling fluid which flows round said cross-flow passages to the trailing edge of the aerofoil to provide cooling thereof.

6. A gas turbine engine having a cooled aerofoil as claimed in any one of the preceding claims.

6. A cooled aerofoil as claimed in claim 5 and in which said second flow passes between an extension of said thn skin and a further contoured region of the trailing edge portion of the contoured suction surface of the aerofoil.

7. A cooled aerofoil substantially as h reinbefore particularly described with reference to the accompanying drawings.

8. A gas turbine engine having a cooled aerofoil as claimed in any one of the preceding claims.

AMENDMENTS TO THE CLAIMS HAVE BEEN FILED AS FOLLOWS 1. A cooled aerofoil for a gas turbine engine comprises a thin skin overlying the contoured surface of the convex suction flank of the aerofoil so as to form therewith a plurality of passages, apertures through which a first flow of coolie fluid may flow into said plurality of passages, at least one cross-flow passage through which the first flow of cooling fluid may flow from said plurality of passages across the section of the aerofoil to the opposed concave pressure flank of the aerofoil, which concave pressure flank has a contoured surface overlaid by a further thin skin having film cooling holes formed therein, said further thin skin overlaying the contoured surface of the concave pressure flank of the aerofoil so as to form therewith a second plurality of passages in communication with said cross-flow passage so that said cooling fluid flows through said second plurality of passages before passing through said film cooling holes.

3. A cooled aerofoil as claimed in any one of the preceding claims and in which there is a second flow of cooling fluid which flows round said cross-flow passages to the trailing edge of the aerofoil to provide cooling thereof.

4. A cooled aerofoil as claimed in claim 3 and in which said second flow passes between an extension of said thin skin and a further contoured region of the trailing edge portion of the contoured suction surface of the aerofoil.

5. A cooled aerofoil substantially as hereinbefore particularly described with reference to the accompanying drawings.