GB1605220A - Blade or vane for a gas turbine engine - Google Patents

Description

(54) A BLADE OR VANE FOR A GAS TURBINE ENGINE (71) We, RoLLs-RoYcE LIMITED.. a British Company, of 65 Buckingham Gate, London SW1E 6AT, ROLLS-ROYCE (1971) LIMITED, a British Company of Norfolk House, St. James's Square, London, S.W.1Y 4JR, do hereby declare the invention, for which we pray that a patent may be granted to us, and the method by which it is to be performed, to be particularly described in and by the following statement: This invention relates to a blade or vane for a gas turbine engine.

Such blades or vanes are often provided with shrouds or platforms which form the outer and/or inner boundaries of the gas flow through the blades or vanes. As the operating temperatures of gas turbine engines have risen it has become necessary to consider the provision of cooling for these shrouds or platforms. One known method of cooling various parts of gas turbine engines which is applicable to shroud cooling, comprises the use of air which is caused to flow through apertures in a plate or other member spaced from that surface of the member to be cooled which is opposed to the gas contacting surface. The cooling fluid, usually air, is caused to flow through the apertures in the form of jets which impinge on this reverse surface and provide cooling.It has, in the past, been a problem to dispose of the cooling fluid once it has performed its task of impingement; it has been proposed to exhaust the air through the trailing edge of the shroud or platform but this involves some aerodynamic losses because of its interference with the normal gas stream of the engine.

The present invention provides a convenient method for disposing of the used cooling fluid.

According to the present invention a blade or vane for a gas turbine engine comprises an aerofoil section and at least one shroud or platform, an apertured impingement plate mounted spaced from that surface of the shroud or platform opposed to its gas contacting surface and adapted to direct cooling fluid to impinge onto said opposed surface of the shroud or platform in the form of a plurality of jets, and passage means within the aerofoil section communicating with the space between the plate and the surface and with apertures extending to the exterior surface of the aerofoil section in the region of the trailing edge so as to conduct, in operation, all of the spent cooling fluid from said space through said apertures and into the main gas flow of the engine.

Said passage means may comprise a duct cast into the structure of the aerofoil section or alternatively it may comprise a fabricated sheet metal tube mounted within the hollow interior of the aerofoil section. Where there are two said shrouds or platforms each mounted at one end of the aerofoil section the impingement plates may be similarly duplicated and may feed both ends of the fabricated or the cast duct.

The invention will now be particularly described merely by way of example with reference to the accompanying drawings in which Figure 1 is a partly broken-away view of a gas turbine engine having vanes in accordance with the invention Figure 2 is a partly broken-away exploded perspective view of one extremity of the nozzle guide vanes of the engine of Figure 1, Figure 3 is a section on the line 3-3 of Figure 2, Figure 4 is a view similar to Figure 2 but of a further embodiment, and Figure 5 is a broken-away exploded perspective view of the cooling arrangement of the trailing edge of the shroud in either of the above embodiments.

In Figure 1 there is shown a gas turbine engine 10 comprising a compressor 11, combustion section 12, turbine section 13 and exhaust nozzle 14. This engine operates in a conventional manner. The casing of the engine is broken away so as to show the downstream end of the combustion chamber 15 and the nozzle guide vanes 16 which serve to direct gases onto the turbine rotor 17. In Figure 2 there is shown a portion of the guide vane 16. Each guide vane comprises an aerofoil section 18 which is mounted between inner and outer shrouds of which only the outer shroud 20 is visible. In order to enable the guide vane to be cooled its interior is made hollow so that within the aerofoil section there is a longitudinally extending hollow 21 in the leading edge and a longitudinally extending hollow 22 which takes up the remainder of the vane.Within the hollow 21 there is mounted a cooling air entry tube 23 which is supplied with cooling air by means not shown and which allows air to impinge on the interior of the leading edge of the vane to provide cooling thereof. Similarly within the trailing edge hollow 22 is mounted a trailing edge tube 24 similarly supplied with cooling air and providing impingement cooling of the vane.

A blanking plate 25 extends over the uppermost extremity of the tubes 23 and 24 to blank off the corresponding areas of the hollows 21 and 22. In order to provide cooling of the shroud 20 an impingement plate 26 is mounted a small distance away from that surface of the shroud 20 opposed to the gas contacting surface. The plate 26 is provided with a plurality of impingement holes 27 through which air from a source not shown is allowed to flow in the form of a plurality of jets to impinge on the surface of the shroud 20 to provide cooling. Once the air has impinged on the shroud it is allowed to flow through that portion of the rearward cavity which is not blanked off by the plate 25 and into a further tube 28 which is mounted behind the tube 24 and extends longitudinally of the cavity.This tube is provided with apertures 29 at its rearward extremity through which the cooling air is allowed to flow and it then passes directly out of the trailing edge of the vane through slots 30. It will thus be seen that the spent cooling air is allowed to flow out of the trailing edge of the vane in a manner which provides fairly small aerodynamic losses.

Figure 4 shows a further embodiment in which there is a similar aerofoil section 31 and shroud 32. The aerofoil section again has a hollow interior divided into forward and rearward cavities 33 and 34 respectively, but in this case the cavity 34 is divided from a further trailing edge cavity 35 by an apertured web 36. Once again a cooling air entry tube 37 is mounted in the cavity 33 and a similar tube 38 is mounted in the cavity 34. These are fed in exactly the same way with cooling air and they act in a similar manner to provide cooling. Again a blanking plate 39 covers the extremities of the tubes 37 and 38 and their associated portions of the cavities 33 and 34.

Once again an impingement plate 40 is mounted just above the shroud 32 and is provided with impingement holes 41 through which air is allowed to flow to impinge on the shroud surface. In this case however, the spent impingement air does not flow into an emit tube such as 28 but is allowed to flow directly into the cavity 35. Here it is joined by the air which flows through the apertured web 36 and the complete flow then leaves the blade through slots 42 in the trailing edge.

It will be seen that this flow re-joins the main gas stream of the engine in a manner giving relatively small aerodynamic losses.

It will be noted that the embodiments described above use impingement cooling over the major area of the shrouds 20 and 32.

However, in both cases there is an area between the edge of the impingement plate and the trailing edge portion of the shroud which is uncooled, and it is possible that it might be necessary to provide cooling for this area.

Figure 5 shows how this cooling may be effected in a convenient manner. Once again an aerofoil portion 45 and shroud 46 are partly shown, with an impingement plate 47 which, when assembled to the shroud, allows cooling of the major part of the shroud. To cool the trailing part of the shroud a plurality of channels 48 are formed in the upper surface of this part of the shroud; these may be formed by chemical etching or other machining, or may be formed when the vane is cast. The channels 48 each extend from a respective entrance chamber 49 toward the rear of the shroud, then extend parallel with the rear edge of the shroud, and turn forward to exit into a discharge plenum 50.The plate 47 is provided with apertures 51 which correspond with the entrance chambers 49 to allow the cooling air to enter the channels, and a cover plate 52 is brazed or otherwise attached to the top surface of the shroud to cover the open face of the channels so that they become separate ducts.

Operation of this system is as follows: the cooling air which is provided for the impingement cooling of the shroud also enters the chambers 49 and flows through the ducts 48 to the plenum 50. The sizes and dispositions of the ducts 48 are chosen to be such as to provide sufficient cooling of the respective part of the shroud, and to arrange that the pressure of the air in the plenum 50 is substantially the same as that of the spent impingement air. The plenum 50 is in flow communication with the passage provided within the aerofoil for discharge of the spent impingement air; this is not shown in Figure 5 but would be similar to the tube 28 or the cavity 35. In this way the shroud trailing edge may be satisfactorily cooled.

It should be noted that there are a considerable number of modifications which could be made to the embodiments described above. Thus in particular, it will be possible to use the arrangement described to cool both shrouds of a vane rather than the single shroud described above. In this case the tube 28 of the cavity 35 would be made to open at both ends of the space between the respective impingement plate and shroud surface. It will also be noted that there could be various different ways of ducting the spent impingement air to the trailing edge holes or slots through which it is allowed to escape.

Although described above as applied to nozzle guide vanes, the invention could clearly be applied to other stators or to rotor blades.

WHAT WE CLAIM IS: 1. A blade or vane for a gas turbine engine comprising an aerofoil section and at least one shroud or platform, an apertured impingement plate mounted spaced from that surface of the shroud or platform opposed to its gas contacting surface and adapted to direct cooling fluid to impinge onto said opposed surface of the shroud or platform in the form of a plurality of jets, and passage means within the aerofoil section communicating with the space between the plate and the surface and with apertures extending to the exterior surface of the aerofoil section in the region of the trailing edge so as to conduct, in operation, all of the spent cooling fluid from said space through said apertures and into the main gas flow of the engine.

2. A blade or vane as claimed in Claim 1 and in which there are two said shrouds or platforms each having one of said plates.

3. A blade or vane as claimed in Claim 1 or Claim 2 and in which said passage means comprises a duct cast into the structure of the aerofoil section.

4. A blade or vane as claimed in Claim 1 or Claim 2 and in which said passage means comprises a fabricated sheet metal tube mounted within the hollow interior of the aerofoil section.

5. A blade or vane as claimed in any preceding claim and in which said impingement plate provides impingement cooling for only the major portion of said shroud or platform, the remainder of the shroud or platform being provided with ducts for the flow of cooling air therein which cool the remainder of the shroud or platform, the ducts also being in flow communication with said passage means.

6. A blade or vane as claimed in any preceding claim and in which said aerofoil section is provided with internal cooling means adapted to receive cooling fluid and to discharge spent cooling fluid into said passage means.

7. A blade or vane as claimed in Claim 7 and in which said internal cooling means comprises an apertured tube or tubes each of which is adapted to receive internally a supply of cooling air and to discharge the cooling air onto the inside of the hollow interior of the aerofoil section so as to cool it.

8. A vane substantially as hereinbefore particularly described with reference to Figures 1, 2 and 3 of the accompanying drawings.

9. A vane substantially as hereinbefore particularly described with reference to Figure 4 of the accompanying drawings.

10. A vane substantially as hereinbefore particularly described with reference to Figure 5 of the accompanying drawings.

11. A gas turbine engine incorporating a blade or vane as claimed in any preceding

Claims (1)

1. claim.