EP1541809B1

EP1541809B1 - Turbine nozzle guide vane and corresponding method of forming

Info

Publication number: EP1541809B1
Application number: EP04257043.2A
Authority: EP
Inventors: Kevin Paul Self; Mark John Simms
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2003-12-12
Filing date: 2004-11-12
Publication date: 2015-03-04
Anticipated expiration: 2024-11-12
Also published as: US20050220619A1; EP1541809A2; EP1541809A3; GB0328952D0

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.
EP1074696 discloses a stator vane having a coolable platform, and EP1162346 discloses a cooling arrangement of a turbine shroud segment.
According to the present invention there is provided a turbine nozzle guide vane for a gas turbine engine according to claim 1.
The invention yet further provides a method of forming turbine nozzle guide vanes for a gas turbine engine according to claim 5.
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 ends 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.

Claims

A turbine nozzle guide vane (10) for a gas turbine engine, the nozzle guide vane (10) including:
a pair of platforms (12,14) with an aerofoil (16) extending therebetween,

a main hollow core provided in each platform,

seal strip slots (20) provided on each end of each platform (12,14) abutment faces (18) provided at each end of each platform (12, 14),

wherein passages (28) are provided extending within the nozzle guide vane (10) from the respective platforms (12,14) to the respective seal strip slots (20), and

characterised in that the passages (28) extend from the main hollow core (32) in the respective platforms (12,14) to the seal strip slots (20) for delivering cooling air to the respective abutment faces (18) of the guide vane (10) the passages and seal strip slots being arranged such that in use cooling air passes through the passages into the seal strip slots and under the seal strip before passing across the abutment face.
A turbine nozzle guide vane according to claim 1, characterised in that the passages (28) are inclined relative to the main hollow core (32).
A turbine nozzle guide vane according to any of the preceding claims, characterised in that a plurality of passages (28) extend to each seal strip slot (20).
A turbine for a gas turbine engine, the turbine including a plurality of nozzle guide vanes (10) arranged in one or more rings, characterised in that the nozzle guide vanes (10) are according to any of the preceding claims.
A method of forming turbine nozzle guide vanes (10) according to claim 1 for a gas turbine engine, the nozzle guide vanes comprising a pair of platforms (12, 14) with an aerofoil extending therebetween and abutment faces provided on each end of each platform (12, 14), the method including:
investment casting metal around a core member, wherein the core member (22) defines openings in the guide vane (10) and wherein the core member defines a main hollow core in each platform and projections (26) on the core member (22) define passages (28),

subsequently removing the core member (22),

providing seal strip slots on each end of each platform, characterised in that the slots are arranged so as to expose ends of the passages (28) such that the passages (28) extend into where the seal strip slots (20) are provided, such that the resulting passages (28) extend from a main hollow core (32) in respective platforms (12,14) to the seal strip slots (20), the passages and seal strip slots being arranged such that in use cooling air passes through the passages into the seal strip slots and under the seal strip before passing across the abutment face.
A method according to claim 5, characterised in that the seal strip slots (20) are machined into the nozzle guide vanes (10) following removal of the core member (22) therefrom, so as to expose ends of said passages (28) in the slots (20).