Nozzle guide vanes

Abstract

A turbine nozzle guide vane with passages leading from a hollow core to respective seal strip slots, to deliver cooling air to abutment faces on each end of the vane.

Description

This invention concerns turbine nozzle guide vanes for gas turbine engines, and a method of forming such nozzle guide vanes.

Turbine nozzle guide vanes for gas turbine engines generally comprise inner and outer platforms with an aerofoil extending therebetween. Such guide vanes are formed as a plurality of segments arranged in one or more rings around an engine. It is necessary for a gap to be left between adjacent guide vanes to allow for manufacturing tolerances and thermal expansion during use. These gaps are conventionally sealed by providing cooperating slots in each guide vane, with a metal seal strip extending in the slots and between the segments.

Nozzle guide vanes are generally air cooled, and passages can be provided in the platforms and aerofoil. It is generally difficult however to cool the abutment faces between adjacent vanes, and particularly due to the provision of the seal strips extending therebetween. Higher engine gas temperatures are generally now being used which make cooling of the nozzle guide vanes increasingly important.

According to the present invention there is provided a turbine nozzle guide vane for a gas turbine engine, the nozzle guide vane including a pair of platforms with an aerofoil extending therebetween, seal strip slots provided on each end of each platform, and passages extending within the nozzle guide vane from the respective platforms to the respective seal strip slots for delivering cooling air to the respective abutment faces of the guide vanes.

The passages preferably extend from a main hollow core in the respective platforms to the seal strip slots.

The passages are preferably inclined relative to the main hollow core. A plurality of passages preferably extend to each seal strip slot.

The invention also provides a turbine for a gas turbine engine, the turbine including a plurality of nozzle guide vanes according to any of the preceding three paragraphs, the nozzle guide vanes being arranged in one or more rings.

The invention yet further provides a method of forming turbine nozzle guide vanes for a gas turbine engine, the method including investment casting metal around a core member, which core member defines openings in the guide vane, subsequently removing the core member, wherein projections on the core member define passages extending into where seal strip slots are provided.

The seal strip slots are preferably machined into the nozzle guide vanes following removal of the core member therefrom, so as to expose ends of said passages in the slots.

An embodiment of the present invention will now be described by way of example only and with reference to the accompanying drawings, in which:

FIG. 1 is a perspective view of a nozzle guide vane according to the invention;

FIG. 2 is a perspective plan view of a core member usable in forming the nozzle guide vane of FIG. 1;

FIG. 3 is a diagrammatic perspective side view of the core member of FIG. 2;

FIG. 4 is a diagrammatic cross sectional side view of part of the guide vane of FIG. 1; and

FIG. 5 is a diagrammatic end view of part of the guide vane of FIG. 1.

FIG. 1 shows a turbine nozzle guide vane 10. The vane 10 has an outer platform 12 and an inner platform 14. An aerofoil 16 extends between the platforms 12, 14. Abutment faces 18 are provided on the end of each of the platforms 12, 14, and seal strip slots 20 are provided in the abutment faces 18.

FIGS. 2 and 3 show a ceramic core member 22 usable in investment casting of the guide vane 10. The core member 22 has a body 24 to define a main hollow core in the guide vane 10, and four inclined projections 26 extending from the body 24 to define passages 28 extending into the seal strip slots 20.

FIGS. 4 and 5 diagrammatically show the nozzle guide vane 10 in use. In FIG. 4 there is shown part of a seal strip 30 locating in the seal strip slot 20. FIG. 4 shows part of an outer platform 12, and above the guide vane 10 as shown in the drawing would be the coolant side at high pressure. Cooling air would be supplied through the main hollow core 32 formed in the body 24 and would then pass through the passages 28 into the seal strip slot 20. The cooling air would generally pass under the seal strip 20 as shown by the arrow, and pass across the abutment face 18 which would face a similar nozzle guide vane 10, to beneath the guide vane 10 as shown, which would be the hot gas side at a lower pressure than the cooling air within the guide vane 10.

In use, the nozzle guide vane 10 would be formed by casting an appropriate metal around the core member 22 in an appropriate shape mould. Following casting the core member 22 would be destroyed, for instance by leaching. The seal strip slots 20 would then be formed by machining until the slot 20 exposes ends of the passages 28. By inclining the projections 26 and hence passages 28, it means that this machining operation will not affect the main hollow core 32 of the guide vane 10.

There is thus described a nozzle guide vane which provides for cooling of the abutment edge and is thus suitable for use at high gas temperatures. No additional manufacturing processes or steps are required in forming such a nozzle guide vane, and therefore such guide vanes can readily be manufactured.

Various modifications may be made without departing from the scope of the invention. For instance, a different number of passages may be provided, and these may be of a different shape.

Whilst endeavouring in the foregoing specification to draw attention to those features of the invention believed to be of particular importance it should be understood that the Applicant claims protection in respect of any patentable feature or combination of features hereinbefore referred to and/or shown in the drawings whether or not particular emphasis has been placed thereon.

Claims

1. A turbine nozzle guide vane for a gas turbine engine, the nozzle guide vane including a pair of platforms with an aerofoil extending therebetween, seal strip slots provided on each end of each platform characterised in that passages are provided extending within the nozzle guide vane from the respective platforms to the respective seal strip slots for delivering cooling air to the respective abutment faces of the guide vanes.

2. A turbine nozzle guide vane according to claim 1, characterised in that the passages extend from a main hollow core in the respective platforms to the seal strip slots.

3. A turbine nozzle guide vane according to claim 2, characterised in that the passages are inclined relative to the main hollow core.

4. A turbine nozzle guide vane according to claim 1, characterised in that a plurality of passages extend to each seal strip slot.

5. A turbine for a gas turbine engine, the turbine including a plurality of nozzle guide vanes arranged in one or more rings, characterised in that the nozzle guide vanes are according to claim 1.

6. A method of forming turbine nozzle guide vanes for a gas turbine engine, the method including investment casting metal around a core member, which core member defines openings in the guide vane, subsequently removing the core member, characterised in that projections on the core member define passages extending into where seal strip slots are provided.

7. A method according to claim 6, characterised in that the seal strip slots are machined into the nozzle guide vanes following removal of the core member therefrom, so as to expose ends of said passages in the slots.