GB2530763A

GB2530763A - A heat shield

Info

Publication number: GB2530763A
Application number: GB1417316.5A
Authority: GB
Inventors: Julian Glyn Balsdon; Peter Lawrence Mason
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2014-10-01
Filing date: 2014-10-01
Publication date: 2016-04-06
Also published as: GB201417316D0

Abstract

A gas turbine engine comprises an aerofoil component 310 having a radially inner platform and a radially outer platform 318 with an aerofoil 312 extending across the main gas path 314 therebetween, and the aerofoil has a fluid pathway therethrough for a flow of cooling air 332. A cavity 324 radially inboard of the inner platform or outboard of outer platform receives a supply of cooling air via a cavity inlet 326. The cavity is fluid communication with the fluid pathway of the aerofoil via an aerofoil inlet 329. A heat shield 330 is located between the cavity inlet 326 and the radially inner or outer platform. The heat shield includes an aperture 348 which is circumferentially aligned with the aerofoil inlet and has a flow area which is approximately the same as or larger than the inlet to the aerofoil. The aerofoil component may be a nozzle guide vane and the heat shield may be a sheet metal insert.

Description

A Heat Shield

Technical Field of Invention

This invention relates to a heat shield for a gas turbine engine. In particular, the invention relates to a heat shield for use with a turbine guide vane.

Background of Invention

With reference to Figure 1, a ducted fan gas turbine engine generally indicated at 10 has a principal and rotational axis X-X. The engine comprises, in axial flow series, an air intake 11, a propulsive fan 12, an intermediate pressure compressor 13, a high-pressure compressor 14, combustion equipment 15, a high-pressure turbine 16, and intermediate pressure turbine 17, a low-pressure turbine 18 and a core engine exhaust nozzle 19. A nacelle 21 generally surrounds the engine 10 and defines the intake 11, a bypass duct 22 and a bypass exhaust nozzle 23.

The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 12 to produce two air flows: a first air flow A into the intermediate pressure compressor 13 and a second air flow B which passes through the bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 13 compresses the air flow A directed into it before delivering that air to the high pressure compressor 14 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 14 is directed into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low-pressure turbines 16, 17, 18 before being exhausted through the nozzle 19 to provide additional propulsive thrust. The high, intermediate and low-pressure turbines respectively drive the high and intermediate pressure compressors 14, 13 and the fan 12 by suitable interconnecting shafts.

The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. It is therefore desirable to operate the turbines at the highest possible temperatures. For any engine cycle compression ratio or bypass ratio, increasing the turbine entry gas temperature produces more specific thrust (e.g. engine thrust per unit of air mass flow). However as turbine entry temperatures increase, the life of an un-cooled turbine falls, necessitating the development of better materials and internal air cooling.

In current engines, the high-pressure turbine gas temperatures are hotter than the melting point of the material of the blades and vanes, necessitating internal air cooling of these airfoil components. During its passage through the engine, the mean temperature of the gas stream decreases as power is extracted. Therefore, the need to cool the static and rotary parts of the engine structure decreases as the gas moves from the high-pressure stage(s), through the intermediate-pressure and low-pressure stages, and towards the exit nozzle.

Figure 2 shows an isometric view of a typical single stage cooled turbine. Cooling air flows are indicated by arrows.

Internal convection and external films are the prime methods of cooling the gas path components -airfoils, platforms, shrouds and shroud segments etc. High-pressure turbine nozzle guide vanes 31 (NGV5) consume the greatest amount of cooling air on high temperature engines. High-pressure blades 32 typically use about half of the NGV flow. The intermediate-pressure and low-pressure stages downstream of the HP turbine use progressively less cooling air.

The high-pressure turbine airfoils are cooled by using high pressure air from the compressor that has by-passed the combustor and is therefore relatively cool compared to the gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while gas temperatures can be in excess of 2100 K. The cooling air from the compressor that is used to cool the hot turbine components is not used fully to extract work from the turbine. Therefore, as extracting coolant flow has an adverse effect on the engine operating efficiency, it is important to use the cooling air effectively.

This invention seeks to provide an improved apparatus for the distribution of cooling air.

Statements of Invention

The present invention provides a gas turbine engine, comprising: an aerofoil component having a radially inner platform and a radially outer platform with an aerofoil extending across the main gas path therebetween, the aerofoil having a fluid pathway therethrough for a flow of cooling air when in use; a cavity radially inboard or outboard of the inner or outer platform for receiving a supply of cooling air via a cavity inlet, the cavity being in fluid communication with the fluid pathway of the aerofoil via an aerofoil inlet; and, a heat shield located between the cavity inlet and the radially inner or outer platform, the heat shield including an aperture which is circumferentially aligned with the aerofoil inlet and has a flow area which is approximately the same as or larger than the inlet to the aerofoil.

Providing a heat shield in this way allows cooling air to be channelled to the NGV without picking up heat load and without cooling the NGV platform.

The gas turbine engine may include an annular array of similar aerofoil components, each aerofoil component having a corresponding heat shield aperture.

The aerofoil component may be a nozzle guide vane. The heat shield may bean annulus.

The heat shield aperture may be a parallelogram.

The heat shield may comprise a plate-like member which extends circumferentially around the engine and a tubular member extending radially inwards from the plate-like member into the aerofoil inlet.

The plate may include platform cooling holes for providing a flow of cooling air to the outboard surface of the platform in use.

The platform cooling holes may be located upstream of the tubular inlet member. The cooling air may flow upstream and downstream. The support flange may include cooling holes.

The heat shield may be snugly received between first and second axially opposing surfaces The first and second axially facing opposing surfaces may be located on separate components of the gas turbine engine.

The second axially opposing surfaces may be provided by a supporting flange which of the aerofoil component.

The heat shield aperture may receive one or more conduits in use.

Description of Drawings

Embodiments of the invention will now be described with the aid of the following drawings of which: Figure 1 shows schematically a longitudinal cross-section through a ducted fan gas turbine engine; Figure 2 shows an isometric view of a typical single stage cooled turbine; Figure 3 shows a schematic representation of a partial cross-section through a Nozzle Guide Vane showing an embodiment of the invention; Figure 4 shows a partial end view of the heat shield shown in Figure 3; and Figure 5 shows an alternative embodiment of the invention.

Detailed Description of Invention

Figures 3 and 4 show, respectively, a longitudinal cross-section and partial end view of an aerofoil and heat shield arrangement of the present invention. Thus, there is shown an aerofoil component in form of a nozzle guide vane, NGV, 310 which finds use in an intermediate pressure turbine stage of a gas turbine engine. The aerofoil 312 of the component is located within the main gas path 314 of the engine and acts to redirect the upstream airflow for the downstream rotatably mounted intermediate pressure turbine blade. The NGV 310 is mounted so as to be stationary relative to the engine casing. The NGV 310 includes a radially inner 316 and radially outer platform 318 which define and bound the main gas path 314 which extends axially along the engine from left to right with respect to Figure 3 and into the page with respect to Figure 4. The outer platform 318 includes radially extending flanges 320 on the outboard side thereof which provide supports for mounting to the engine casing 322 via bird mouth attachments as known in the art. The support flanges 320 extend in a circumferential line around the outboard surface at a constant axial position.

Upstream of the NGV 310 just out of view to the left of Figure 3 is located a seal segment 35 of a turbine blade 32 of the previous stage, which in the described embodiment, is a high pressure turbine as shown in Figure 2. Downstream is a seal segment of the intermediate pressure turbine blade.

The engine casing 322 resides outboard of the platform 318 and defines a cavity 324 therebetween for receiving cooling air from an upstream air source which is typically an appropriate stage of the gas turbine compressor, as is known in the art. The air is delivered to the cavity 324 by means of an inlet 326 which is attached to a conduit 328 which is fluidically connected to the compressed air source.

As can best be seen from Figure 4, the aerofoil 312 component is substantially hollow to allow cooling air to flow therethrough. Thus, the radially outer end of the NGV 310 has a cooling air inlet 329 in the form of the substantially open end of the aerofoil 312. The open end is as large as can be accommodated and therefore predominantly a continuation of the internal surface of the NGV 310 wall with any necessary thickening required forjoining the vane to the platform. The cooling air may be used in conjunction with internal turbulating features or to provide film cooling air on the exterior of the aerofoil 312 as is well known in the art.

The cooling air cavity 324 on the outboard side of the platform 318 includes a heat shield 330 which segregates the platform 318 and cooling air inlet 328 such that the cooling air 332 cannot impinge directly onto the NGV platform 318. This reduces the heat pick up of the cooling air 332 from the platform 318 and helps reduce concentrated cooling of the platform 318 in the area radially inwards of the conduit inlet 328. A further advantage of using the heat shield 330 is that it allows the circumferential distribution of cooling air inlet manifolds to be non-uniform around the casing whilst still allowing distribution of the cooling air limited detriment to the cooling performance of the NGVS.

The heat shield 330 of the described embodiment is fabricated as a sheet metal annular insert. The heat shield 330 includes an upstream end 334 and a downstream end 336 with a

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platform section 338 therebetween. The general shape of the heat shield 330 in section as shown in Figure 3, corresponds closely to the internal profile of the NGV platform 318, with sufficient clearance to allow some relative lateral movement without contact and the associated wear and heat transfer from the platform 318. The upstream end 334 includes a contact portion 340 which abuts a portion of wall 320 which extends radially away from the main gas path 314 and faces axially downstream to oppose the upstream contacting portion 340. The upstream contacting portion 340 includes a radial outturn 342 to provide a suitable contacting surface which allows a compressive sealing abutment with the engine casing 322 wall. The downstream end 336 includes a radial outturn 344 which is followed by an axial back turn such that the downstream free end 346 of the heat shield 330 projects upstream and is suspended over the main part of the heat shield 330. The radial outturn 344 provides a contacting surface which faces axially downstream and abuts a portion of the supporting flange of the NGV 310. The back turned portion provides a contacting surface which faces radially outwards and abuts the radial inner side of the engine casing 320 birds mouth feature.

The combination of the upstream and downstieam contacting portions provides corresponding abutting sealing surfaces at the fore and aft ends of thereof. The heat shield 330 is sized marginally larger than the cooling air cavity 324 such that is compressed slightly when the engine is assembled. Having a compress fit such as this helps maintain the seal at the peripheries of the heat shield 330 and aids the channelling of the cooling air 332.

As can be seen best in Figure 3, the heat shield 330 includes a plurality of cooling air apertures 348 which reside radially outwards and correspond to open ends of the NGVs.

Thus, for each NGV 310 there is an opening 348 in the heat shield 330. Each opening 348 in the heat shield 330 is sized to provide a minimum amount of resistance to the cooling air 332 as it passes into the NGV 310 and is therefore larger than the corresponding opening 348 in each NGV 310. In the described embodiment, the heat shield openings 348 are parallelograms in shape having two sides which extend around circumferential lines in planes normal to the principal axis of the engine, and two sides which are angled to roughly correspond to the chord of the pressure surface of the vanes. It will be appreciated that the aperture 329 may be any suitable shape which provides minimal resistance to the flow entering the aerofoil.

The heat shield 330 includes incidental ancillary features such as anti-rotation tangs 350 which engage with corresponding features of the NGV.

The heat shield 330 may be fabricated from any suitable sheet material such as nickel based alloy but it will be apparent that the choice of material will be application specific. The fabrication of the heat shield 330 can be carried out using well known methods known the art.

To assemble the heat shield 330 into the engine, it is placed over the NGV 310 from an upstream direction prior to the bird mouth attachments being inserted into the corresponding features on the interior of the engine casing 320.

Figure 5 shows an alternative embodiment on the invention in which the heat shield 530 includes a tubular component 552 which extends radially inwards in to the hollow interior of the NGV 510. This arrangement provides a more direct way to channel the cooling air 532 into the vane 510 with no leakage to the space which separates the heat shield 530 and platform 518.

In either embodiment, the heat shield 530 may incorporate apertures 554 to allow cooling air 556 to be metered onto the platform 518 to allow for some controlled flow to cool the platform. In the described embodiment, the cooling air aperture 554 is placed towards an upstream end of the heat shield 530 such that cooling air can enter the space and travel downstream to an exhaust aperture located towards the trailing edge of the platform 518 where the pressure is lower.

Also contemplated in this embodiment is the provision of cooling air upstream to feed the axially adjacent upstream seal segment.

Claims

Claims: 1. A gas turbine engine (10), comprising: an aerofoil component (310) having a radially inner platform (316) and a radially outer (318) platform with an aerofoil (312) extending across the main gas path (314) therebetween, the aerofoil having a fluid pathway therethrough for a flow of cooling air (332) when in use; a cavity (324) radially inboard or outboard of the inner or outer platform for receiving a supply of cooling air via a cavity inlet (326), the cavity being in fluid communication with the fluid pathway of the aerofoil via an aerofoil inlet (329); and, a heat shield (330) located between the cavity inlet (326) and the radially inner or outer platform, the heat shield including an aperture (348) which is circumferentially aligned with the aerofoil inlet and has a flow area which is approximately the same as or larger than the inlet to the aerofoil.
2. A gas turbine engine as claimed in claim 1, wherein the gas turbine engine includes an annular array of similar aerofoil components, each aerofoil component having a corresponding heat shield aperture.
3. A gas turbine as claimed in claims 1 or 2, wherein the aerofoil component is a nozzle guide vane.
4. A gas turbine engine as claimed in any preceding claim, wherein the heat shield is an annulus.
5. A gas turbine engine as claimed in any preceding claim, wherein the heat shield aperture is a parallelogram.
6. A gas turbine engine as claimed in any preceding claim, wherein the heat shield comprises a plate-like member which extends circumferentially around the engine and a tubular member extending radially inwards from the plate-like member into the aerofoil inlet.
7. A gas turbine engine as claimed in any preceding claim, wherein the plate includes platform cooling holes for providing a flow of cooling air to the outboard surface of the platform in use.
8. A gas turbine engine as claimed in any preceding claim, wherein the platform cooling holes are located upstream of the tubular inlet member.
9. A gas turbine engine as claimed in any preceding claim, wherein the cooling air flows upstream and downstream.
10. A gas turbine engine as claimed in any preceding claim, wherein the support flange includes cooling holes.
11. A gas turbine engine as claimed in any preceding claim, wherein the heat shield is snugly received between first and second axially opposing surfaces
12. A gas turbine engine as claimed in any preceding claim, wherein the first and second axially facing opposing surfaces are located on separate components of the gas turbine engine.
13. A gas turbine engine as claimed in any preceding claim, wherein the second axially opposing surfaces is provided by a supporting flange which of the aerofoil component.
14. A gas turbine engine as claimed in any preceding claim, wherein the heat shield aperture receives one or more conduits in use.