GB2235253A

GB2235253A - Ceramic guide vane for gas turbine engine

Info

Publication number: GB2235253A
Application number: GB8918683A
Authority: GB
Inventors: Paul Robert Hayton
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1989-08-16
Filing date: 1989-08-16
Publication date: 1991-02-27
Also published as: GB8918683D0

Abstract

Ceramic nozzle guide vanes 22 for a gas turbine engine are each provided with an integral ceramic shroud member 28 having upstream and downstream sections 30, 32 which are in contact with the engine turbine casing solely at upstream and downstream extremities thereof. Cooling air channels are provided between the shroud members 28 and the turbine casing. Split sealing rings 66, 68 are provided circumferentially of the shroud members to seal any gaps between adjacent shroud members. A rod 52, having a channel 54 is provided radially through each vane 22 and its shroud member 28 to allow for the passage of cooling air between a region outside the turbine casing and a region radially inwards of the vane. Ring seals 44, are provided at the radially inner ends of the vanes. <IMAGE>

Description

Improvements in or relating to nozzle guide vanes for gas turbine engines This invention concerns improvements in or relating to nozzle guide vanes for gas turbine engines, and in particular is concerned with the amelioration of air leakage between vane segments and the minimisation of cooling air needed to cool t:ie vanes.

In the operation of a gas turbine engine air at atmospheric pressure is initially compressed by a compressor and delivered to a combustion stage. In the combustion stage heat is added to the air leaving the compressor by adding fuel to the air and burning it.

The gas flow resulting fym combustion of fuel in the combustion stage then expands through a turbine, delivering up some of its energy to drive the turbine and produce mechanical power, the remainder, on discharge to atmosphere, providing a propulsive jet in the case of gas turbine jet engine, for example.

In order to produce a driving torque, the turbine consists of one or more stages, each employing one row of stationary nozzle guide vanes and one row of moving blades mounted on or integral with a disc. The nozzle guide vanes are aerodynamically designed to direct incoming gas form the combustion stage out the turbine blades and thereby transfer kinetic energy to the blades.

Gases entering the turbine typically have an entry temperature form 8500 to at least 17000C. Since the efficiency and power output of the turbine are functions of the entry temperature of the incoming gases there is a trend in gas turbine engine technology to increase the gas temperature. A consequence of this is that the materials of which the blades and vanes are made assume ever-increasing importance with a view to resisting the effects of elevated temperature.

Historically, nozzle guide vanes have been made of metals such as high temperature steels and , latterly, nickel alloys, and it has been found necessary to provide internal cooling passages within the vanes in order to prevent over-heating and melting. It has been found that resistance to even higher gas entry temperatures may be imparted to the vanes by making the vanes entirely of ceramic, thus doing away with the need for internal cooling passages.

However, it is still necessary to keep the turbine casing cool, to which nozzle guide vanes are attached, by preventing, as far as possible, hot gases from impinging on the turbine casing.

It is an object of the invention to provide nozzle guide vanes which minimise heating of the turbine casing and which minimise the amount of cooling air needed to cool the vanes.

For convenience, a brief description of a ducted fan gas turbine engine to which the invention will be applied is now provided with reference to Figure 1 of the drawings.

Figure 1 is a sectioned side view of a ducted fan gas turbine engine, indicated generally at 10, comprising, in axial fluid flow series, an ducted fan 11, a compressor section 12, combustion equipment 13, a turbine section 14, and a propulsion nozzle 15. The turbine section 14 is interconnected with the compressor section 12 and the ducted fan 11 by means of axial shafts 17, and includes a circumferential array 18 of ceramic nozzle guide vanes located immediately downstream of the combustion equipment 13.

The engine 10 functions in the conventional manner with air compressed by the ducted fan 11 being split into two flows. The first flow is directed through outlet guide vanes 16 to provide propulsive thrust, whilst the second flow is directed into the compressor section 12 where it is compressed still further. The compressed air is then directed into the combustion equipment 13 where it is mixed with fuel and the mixture combusted to provide hot gases which expanded through the turbine section 14 and are exhausted at atmosphere through the nozzle 15 to provide propulsive thrust.

According to the present invention there is provided a nozzle guide vane for a gas turbine engine provided with a turbine casing characterised in that the vane is provided with an integral ceramic shroud member at that end of the vane radially outermost with respect to the longitudipal axis of the engine.

The invention will now be described by way of example with reference to Figures 2-5 of the accompanying diagrammatic non-scale drawings in which, Figure 2 is a sectioned side view of part of the turbine section of the engine of Figure 1 showing a section through a nozzle guide vane; Figure 3 is a view of the section of Figure 2 taken in the direction of arrow III; and Figures 4 and 5 are detail sections of edges of adjacent nozzle guide vanes.

"Radially" in this specification will be understood to mean the direction substantially at right angles to the longitudinal axis 9 of the engine section shown in Figure 1.

Referring to Figure 2, and where necessary, Figure 3, there is shown a portion of a-cylindrical turbine casing 20 and, within the casing, a.nozzle guide vane 22 of the array 18. The nozzle guide vane 22 comprises an aerofoil section 24, and radially inner platform 26, and an integral shroud member 28 located radially outwardly of the aerofoil section. The shroud member 28 extends longitudinally upstream and downstream of the aerofoil section 24 as upstream and downstream sections 30, 32 respectively.

The upstream section 30 of the shroud member 28 is provided at its upstream end with a radially outwardly extending lip 34.

Likewise, the downstream end of downstream section 32 of the shroud member 28 is provided with a radially outwardly extending lip 36. Lips 34, 36 about the inner face of the turbine casing 20 and are held in position by respective anti-rotation pins 38, 40.

The inner platform 26 of the nozzle guide vane 22 is supported against a seal ring interstage 42 by ring seals 44.

A channel 46 extends radially through the platform 26, aerofoil section 24, and shroud member 28 of the nozzle guide vane 22. A second channel 48, in alignment with channel 46, extends through the turbine casing 20. A third channel 50, also in alignment with channel 46, extends through the seal ring interstage 42.

A rod 52 having axial passage 54 extending there through from end to end is located through channels 46, 48 and 50, and is provided with a flange 56 at one end overlapping the radially outer surface of the turbine casing 20. The flange 56 is held to the casing 20 by welds 58, and sealed by an interference fit. Hence, the hollow rod 52 provides means by which cooling air can pass between a region radially inwards of the seal ring interstage 42 and a region external of the turbine casing 20 and lying between the turbine casing and a surrounding impingement plate 60.

Because of the radially outwardly extending lips 34, 36 of the shroud member 28 abutting the casing 20, annular cavities 62, 64 are formed between the upstream and downstream sections 30, 32 respectively of the shroud member and the casing. The cavities 62, 64 enable cooling air to pass between the shroud member 28 and the turbine casing 20.

However, each nozzle guide vane 22, together with its integral shroud member 28, is a segmental part of the circumferential array 18 of nozzle guide vanes. When the engine is operating it is possible for hot gases from the combustion equipment 13 to leak between the vane segments into the annular cavities 62 and 64, thus reducing the cooling effect of the cooling air within the cavities. This problem is overcome by the provision of split rings 66, 68 within respective cavities 62, 64.

The split rings 66, 68 are in compression and in face-to-face contact with the radially outer surface of the shroud member, thus covering the gaps between the vane segments and preventing hot gases from leaking into the cavities 62, 64. The radially outer surface of the upstream section 30 of the shroud member recessed to receive the upstream split ring 66. The downstream section 32 of the shroud member 28 is not recessed to receive the downstream split ring 68, but the turbine casing 20 is provided with inwardly directed upstream and downstream radial flanges 70, 72 respectively which abut and seal opposed edges of ring 68.

For successful operation of sealing rings 66 and 68, the air pressure in the chambers 62 and 64 must be, on average, greater than that at the face between the ring seals and the shrouds/nozzle outer face. This may be difficult to achieve and, if achieved, may exert unacceptable bending loads on the nozzle. In this case, a nozzle-to-nozzle sealing arrangement would be used wherein a thin slot 74 is produced in the shroud edge, and a thin high temperature strip 76 is placed in the thin slot, as shown in Figures 4 and 5. To prevent load transfers between nozzles the slot gap is greater than the thickness of the strip 76 so as to permit relative movement between the nozzles (Figures 4 and 5).

In the invention, the integration of the ceramic shroud member with the ceramic nozzle guide vane keeps the maximum amount of metallic engine structure clear of the lost gas stream so as to minimize the amount of cooling air necessary to cool the structure. The replacement of metallic components in this manner by ceramic components which can withstand high temperatures reduces the need to use air to cool the components with the results that engine efficiency is improved. Furthermore, since the only contact of the nozzle guide vane 22 with the turbine casing 20 is by means of relatively small contact areas at the lips 34, 36 of the upstream and downstream sections 30, 32 respectively of the integral shroud number 28, there is minimum heat transfer form the guide vane to the casing. It will also be observed that thermal expansion of the guide vane 22 in the radial direction will be accommodated by cantilever bending of the shroud member 28 between the lips 34, 36.

Modifications within the scope of the invention may be envisaged in which the shroud member 28 does not extend both upstream and downstream of the vane, but in one direction only. Furthermore, the shroud member 28 may be supported against the casing 20 at locations other than the upstream and downstream extremities 34,36.

Types of seals other than the upstream and downstream split rings 66, 68 may be envisaged. For example, sealing between adjacent shroud members may be by an interlocking arrangement of the shrouds. However, this may diminish independence of the vanes and permit an undesirable stacking up of vane loads under some circumstances.

Claims

Claims

1A ceramic nozzle guide vane for a gas turbine engine provided with a turbine casing characterised in that the vane is provided with an integral ceramic shroud member at that end of the vane radially outermost with respect to the longitudinal axis of the engine.
2 A nozzle guide vane is claimed in claim 1 characterised in that the shroud member extends upstream and downstream of the vane.
3 A nozzle guide vane as claimed in claim 2 characterised in that the vane is in contact with the engine turbine casing solely at upstream and downstream extremities of the shroud member.
4 A nozzle guide vane as claimed in any preceding claim wherein there is provided sealing means to seal a gap between the shroud member and a shroud member of a adjacent vane thereby to prevent the passage of hot engine gases to a region between the shroud member and the turbine casing.
5 A nozzle guide vane as claimed in claim 4 wherein the sealing means is provided by at least one split.

ring circumferentially enclosing at least one portion of the shroud member.
6 A nozzle guide vane as claimed in claim 4 wherein upstream and downstream edges of the or each split ring are in sealing abutment with radially extending portions of the shroud member.
7 A nozzle guide vane is claimed in claim 4 wherein upstream and downstream edges of the or each split ring are in sealing abutment with radially extending portions of the turbine casing.
8 A nozzle guide vane as claimed in any preceding claim wherein the radially innermost portion of the vane is provided with a platform sealing against a seal ring interstage portion of the engine.
9 A nozzle guide vane as claimed in claim 8 wherein there is provided channel means radially through the vane and shroud member to provide for the passage of cooling air between a region radially outwards of the turbine casing and a region radially inwards of the seal ring interstage.
10 A nozzle guide vane as claimed in any preceding claim wherein there is provided at least one cooling air channel between the shroud member and the turbine casing.
11 A nozzle guide vane as claimed in any preceding claim wherein there is provided at least one anti-rotation pin fastening the shroud member to the turbine casing.
12 A nozzle guide vane substantially as herein described with reference to Figures 2 to 5 of the accomDanvinq drawinqs.