US3384346A - Aerofoil shaped blade for a fluid flow machine such as a gas turbine engine - Google Patents

Description

May 21, 1968 G. A. HALLS ET AL 3,334,346

AEROFOIL SHAPED BLADE FOR A FLUID FLOW MACHINE SUCH AS A GAS TURBINE ENGINE Filed Jan. 19, 1967 L. .f v

United States Patent O ABSTRACT F THE DISCLOSURE An aerofoil-shaped blade having internal and external walls separated from each other by a space. The interior of the internal wall is adapted to receive cooling fluid and part of this cooling fluid passes through a multiplicity of holes in the internal wall and impinges on the interior of the external wall. The remainder of the cooling uid passes across the space between the two walls of the blade directly to the exterior of the blade without mixing with the cooling fluid in the space.

This invention concerns an aerofoil-shaped blade adapted for use in a uid flow machine such as a gas turbine engine.

The term blade is used in this specification in a broad sense as including for example nozzle guide vanes.

According to the present invention, there is provided an aerofoil-shaped blade adapted for use in a fluid tlow machine such as a gas turbine engine and having an internal Wall which is surrounded by an external Wall and is separated therefrom by a space which extends adjacent both the opposite sides of the blade, the internal wall dening a chamber which is adapted to receive a cooling fluid, at least a major portion of the internal wall being provided with a multiplicity of holes through which the cooling uid may pass from said chamber to said space so as to ow over and cool the interior of the external wall, the external wall having at least one aperture enabling the cooling Huid to escape from the space, said blade being provided with at least one passage through which a portion of the cooling fluid entering the chamber may pass across said space directly to the exterior of the blade without mixing with the cooling fluid in the space.

It will be appreciated that the present invention permits substantially the whole of the interior of the external wall of the blade to be cooled by jets of cooling fluid from the said chamber and thus to be more effectively cooled than hitherto. Moreover, the said multiplicity of holes may be so arranged as to ensure adequate cooling of the said interior of the external wall notwithstanding mal-distribution of the air ow in the said space. Thus by providing the holes in the right places, it is not necessary to use batlies in the said space to ensure that cooling uid is directed to all parts thereof, Moreover, the disposition of the holes may be such as to compensate for local hot spots, while the provision of these holes over at least a major portion of the internal wall makes the disposition of the latter with respect to the external wall less critical than would otherwise be the case.

The said holes preferably extend throughout at least a rnajor portion of both the radial length and the axial length of the internalwall.

The said holes may include holes extending throughout the whole radial length of the portion of the internal wall adjacent the leading edge of the blade.

Preferably few or no holes are provided adjacent the radially inner and outer ends of the internal wall.

The radial length of the blade at its upstream en-d may 3,384,346 Patented May 21, 1968 be substantially greater than its radial length at its downstream end. l

The said space may extend to the trailing edge of the blade, the external wall being provided with a said aperture or apertures at the said trailing edge.

The invention also comprises an annular combustion chamber for a gas turbine engine `having a plurality of angularly spaced apart nozzle guide vanes therein constituted by blades as set forth above.

The radial length of the combustion chamber may be substantially greater centrally of its ends than at either of its ends.

The invention is illustrated, merely by way of example, in the accompanying drawings, in which:

FIGURE 1 is a diagrammatic view, partly in section, of a gas turbine engine having nozzle guide vanes in accordance with the present invention,

FIGURE 2 is an enlarged sectional View of such a nozzle guide vane, and

FIGURE 3 is a section taken on the line 3-3 of FIG- URE 2.

In FIGURE l there is shown a gas turbine engine 10 having a compressor 11, a combustion chamber 12, and a turbine 13 which drives the compressor 11, the turbine exhaust gases being directed to atmosphere through an exhaust pipe 14.

The combustion chamber 12 is an annular combustion chamber in the downstream 'half of which there are mounted a plurality of angularly spaced apart nozzle guide vanes 15.

Each of the nozzle guide vanes 15 has an internal wall 16 which defines a chamber 17. The chamber 17 is supplied at its radially outer end with part of the air compressed by the compressor 11, this air being employed as dilution air for the dilution of the products of combustion so that their temperature is acceptable to the turbine 13. In order to simplify FIGURE l, however, the manner in which the dilution air enters the chamber 17 is not shown.

The internal wall 16 of each nozzle guide vane 15 is surrounded by an external wall 20 thereof. The walls 16, 20 are separated by a space 21 which extends adjacent both the opposite sides 22, 23 of the nozzle guide vane and extends to the trailing edge 24 thereof.

A major portion of the internal wall 16 is provided with a multiplicity of holes 25 through which the dilution air may pass from the chamber 17 to the space 21 so as to ow over and cool the interior of the external wall 20. The air passing through the holes 25 irnpinges on the external Wall 20 in the form of jets. The external wall 20 is provided at the trailing edge 24 with one or more apertures 26 through which the dilution air may escape from the space 21.

As will be seen, the holes 25 extend throughout a major portion of both the radial length and the axial length of the internal wall 16. The holes 25 also include a row of holes 25a which extend throughout the whole radial length of the portion of the internal wall 16 adjacent the leading edge 30 of the nozzle guide vane. Thus, there is adequate cooling of the said leading edge. Very few of the holes 25 however are provided adjacent the radially inner and outer ends of the internal wall 16.

Each nozzle guide vane 15 is provided with a plurality of radially spaced apart passages 31 through which a portion of the dilution air which has entered the chamber 17 rnay pass across the space 21 directly to the interior of the nozzle guide vane without mixing with the dilution air in the space 21.

The downstream end of the internal wall 16 is spaced well Iaway from the trailing edge 24, ow through the portion of the space 21 downstream of the internal wall 16 :being divided into three portions by internal battles 32.

3. ...The radial length of .the combustion chamber 12 is substantially greater centrally of its ends than at either of its ends. As a result, the nozzle guide vanes 15 have to be formed las shown with a radial length at their upstream ends which is lsubstantially greater than their radial length at their downstream ends. Th-us, the radially innermost portion lof the space 21 will, in operation, form .a low pressure region which, but for the fact that the holes 25 are provided throughout a major portion of the internal wall 16, would result in inadequate cooling of some parts of the interior of the external Wall 20.

We claim:

1. An aerofoil-shaped blade yadapted for use in a fluid ow machine such as 4a gas turbine engine, comprising an internal Wall and an external Wall, the said internal wall being surrounded by said external wall and being sepa rated entirely therefrom by a space which extends adjacent lboth the opposite sides of the blade, a chamber deiined-by said internal Wall, said chamber being adapted to receive a cooling fluid, a multiplicity .of holes being .provided on at least a major portion of the internal Wall through which the cooling uid may pass from said chamber to said space so as to How over and cool the entire interior of the external wall, at least yone aperture being provided in the external wall enabling the cooling fluid to escape from the space, at least one passage in said blade extending directly lfrom the chamber within said internal wall across said space to the exterior of said external wall through which a portion of the cooling fluid entering the chamber may pass across said space directly to the exterior of the blade without mixing with the cooling fluid in the space.

2. A blade as claimed .in claim 1 in which the said holes extend throughout at least a major portion of both the radial length and the axial length of the internal Wall.

3. A [blade as claimed in claim 1 in which the said holes include holes extending throughout the Whole radial length of the portion of the internal Wall adjacent the leading edge of the blade.

4. A blade as claimed in claim 1 in which a limited number of holes are provided adjacent the radially inner and outer ends of the internal Wall.

5. A blade as claimed in claim 1 in which the radial length of the blade 'at its upstream end is substantially greater than its radial length -at its downstream end.

6. A blade as claimed in claim 1 in which the said space extends to the trailing edge of the blade, the external wall being provided with at least one said aperture at the said trailing edge.

7. An annular combustion cham-ber for a gas turbine engine having a plurality of angularly spaced apart nozzle guide vfanes therein constituted by blades as claimed in claim 1.

8. A combustion chamber as claimed in claim 7 in which the radial length of the combustion chamber is substantially greater centrally of its ends than at either of its ends.

References Cited UNITED STATES PATENTS 2,647,368 8/1953 Triebbnigg et al. 253-3915 2,873,944 2/1959 Wiese et al 253-3915 2,879,028 3/1959 Stalker 253-3915 3,032,314 5/1962 David 253-3915 3,246,469 4/1966 Moore 253-391 3,299,632 l/l967 Wilde et al. 253-39.l 3,301,527 l/l967 Kercher 253-39.1

EVERETTE A. POWELL, JR., Primary Examiner.

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