GB2467350A

GB2467350A - Cooling and sealing in gas turbine engine turbine stage

Info

Publication number: GB2467350A
Application number: GB0901547A
Authority: GB
Inventors: Roderick Miles Townes
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2009-02-02
Filing date: 2009-02-02
Publication date: 2010-08-04
Also published as: GB0901547D0

Abstract

A turbine stage of a gas turbine engine has a row of rotor blades 102 mounted at their ends to a rim of a rotor disc 108, and a row of nozzle guide vanes. The blades and vanes have internal conduits 104 for carrying a cooling gas flow. The blades have respective platforms 109 having front and rear edges which form corresponding front and rear annular gaps with facing stationary portions 110, 111 (which may comprise nozzle guide vane platforms) of the engine. The stage is adapted to produce a sealing gas flow B which issues from the gaps into the annular passage to prevent the working gas leaking in the opposite direction. At least a portion A of the cooling gas flow is reused and may pass through an opening 112 to form at least a portion of the sealing gas flow. Respective front and rear openings may be provided in the blades to provide sealing airflow at both front and rear annular gaps.

Description

TURBINE STAGE

The present invention relates to a turbine stage of a gas turbine engine.

The performance of gas turbine engines, whether measured in terms of efficiency or specific output, is improved by increasing the turbine gas temperature. In modern engines, the high pressure (HP) working 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. Nonetheless, in some engines, the intermediate pressure (IP) and low pressure (LP) turbines are also internally cooled.

Figure 1 shows an isometric view of a typical HP stage of a cooled turbine. Cooling air flows are indicated by arrows.

Internal convection and external films are the prime methods of cooling the airfoils. HP turbine nozzle guide vanes 1 (NGV5) consume the greatest amount of cooling air on high temperature engines. HP blades 2 typically use about half of the NGV flow. The IP and LP stages downstream of the HP turbine use progressively less cooling air.

The HP 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 working gas temperature. Typical cooling air temperatures are between 800 and 1000 K, while working gas temperatures can be in excess of 2100 K. Cooling air is carried along internal conduits within the airfoils and exits through a large number of cooling holes formed in the surfaces of the airfoils.

The cooling air bled 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 minimise the use of cooling air and maximise its effectiveness.

Figure 2 shows a cross-sectional view of a blade 2 for a stage similar to the stage of Figure 1. Cooling air (indicated by arrows A) bled from the compressor enters the root 3 of the blade, and flows along an internal conduit 4 (indicated by dotted lines) which extends from the root.

Cooling air that does not exit through cooling holes 5 formed in the surfaces of the blade exits the blade at an opening or openings in a shroud 6 formed at the tip of the blade.

The blades 2 are mounted to the rim of a disc 8, the rim being positioned in a disc rim cavity 7 which is bounded at its radially outer side by a blade platform 9 and facing front 10 and rear 11 NGV platforms. A front annular gap is formed with the blade platform 9 on one side of the gap and the front NGV platform 10 and a radially inner sealing flange 14 on the other side of the gap.

Likewise, a rear annular gap is formed with the blade platform 9 on one side of the gap and the rear NGV platform 11 and another radially inner sealing flange 15 on the other side of the gap. A flow of sealing air (indicated by arrows B), also bled from the compressor, issues through the gaps from the disc rim cavity to prevent the high temperature working gas leaking into the cavity.

A first aspect of the invention provides a turbine stage of a gas turbine engine, the stage having: an annular passage for conveying a flow of a working gas, a rotor disc, a row of rotor blades mounted at their ends to a rim of the rotor disc, and a corresponding row of nozzle guide vanes; wherein: the blades and vanes extend across the annular passage and, in use, extract power from the working gas, the blades and vanes having internal conduits for carrying a cooling gas flow, the blades have respective platforms adjacent said ends which define a boundary of the annular passage, front and rear edges of the platforms forming corresponding front and rear annular gaps with facing stationary portions of the engine, and the stage is adapted to produce a sealing gas flow which issues from the gaps into the annular passage to prevent the working gas leaking in the opposite direction; wherein at least a portion of the cooling gas flow is reused to form at least a portion of the sealing gas flow.

Preferably the stage is a high or intermediate pressure turbine stage.

By reusing the cooling gas flow in this way, a proportion of the sealing gas flow which is needed for a conventional stage can be replaced. This can lead to a drop in overall compressor bleed flow, and an increase in engine efficiency.

Preferably, the internal conduits of the rotor blades route the cooling gas carried therein towards the rim of the rotor disc, the blades having respective openings at said ends through which cooling gas escapes from the internal conduits and forms at least a portion of the sealing gas flow. The escaped cooling gas flow can form at least a portion of the sealing gas flow through the rear annular gap and/or at least a portion of the sealing gas flow through the front annular gap.

Having cooling air entering the annular passage (in the sealing gas flow) closer to the hub of the engine can help to reduce aerodynamic spoiling losses. In contrast, in a conventional stage where a greater proportion of the cooling air enters the annular passage from the tip of the blade, the losses can be higher.

Additionally or alternatively, the stage may be adapted so that cooling gas flow which has passed through the internal conduits of the nozzle guide vanes is routed to form at least a portion of the sealing gas flow.

A second aspect of the invention provides a gas turbine engine having the turbine stage of the first aspect.

A third aspect of the invention provides a rotor blade for use in the turbine stage of the first aspect.

A fourth aspect of the invention provides a nozzle guide vane for use in the turbine stage of the first aspect.

Embodiments of the invention will now be described by way of example with reference to the accompanying drawings in which: Figure 1 shows an isometric view of a typical HP stage of a cooled turbine; Figure 2 shows a cross-sectional view of a blade for a stage similar to the stage of Figure 1; Figure 3 shows a cross-sectional view of a blade for a turbine stage according to a first embodiment of the invention; and Figure 4 shows a cross-sectional view of a blade for a turbine stage according to a second embodiment of the invention.

Figure 3 shows a cross-sectional view of a blade for a turbine stage according to a first embodiment of the invention.

The blade 102 is mounted to the rim of a rotor disc 108 and extends across an annular passage through which working gas is conveyed. Cooling air (indicated by arrows A) bled from the compressor enters the root 103 of a blade 102, and flows along an internal conduit 104 (indicated by dotted lines) which has an "up" leg extending from the root. A portion of the cooling air continues from the "up" leg into a shroud 106 at the tip, but the remaining cooling air follows the conduit around a 180° bend and along a "down" leg that returns the cooling air towards the root.

Cooling air that does not exit through cooling holes (not shown) formed in the surfaces of the blade exits the blade at an opening 112 at the rear side of the blade beneath a platform 109 which defines a boundary of the annular passage across which the blade extends. A plate or weld 113 at the base of the "down" leg prevents the cooling air from exiting at the root of the blade. Alternatively, by use of a suitably-shaped core during casting of the blade, the conduit can be formed during casting such that it does not extend to the base of the blade, thereby avoiding the need for a further forming operation.

The rim of the disc 108 is positioned in a disc rim cavity 107 which is bounded at its radially outer side by the blade platform 109 and facing front 110 and rear 111 NGV platforms. A front annular gap is formed with the blade platform 109 on one side of the gap and the front NGV platform 110 and a radially inner sealing flange 114 on the other side of the gap. Likewise, a rear annular gap is formed with the blade platform 9 on one side of the gap and the rear NGV platform 111 and another radially inner sealing flange 115 on the other side of the gap. The NGV platforms and sealing flanges are stationary portions of the engine. A flow of sealing air (indicated by arrows B), which is also bled from the compressor, issues through the gaps from the disc rim cavity. However, the flow through the rear gap is reduced (as indicated by the dashed arrow B) relative to conventional levels as the sealing air flow for this gap is supplemented by the cooling air exiting opening 112. This reused cooling air in combination with the reduced flow B through the rear gap is sufficient to create a sealing flow from the gap that prevents the working gas leaking into the cavity.

Advantageously, in this way the overall amount of air that needs to be bled from the compressor can be reduced, which improves engine efficiency. Alternatively, the overall amount of air bled from the compressor can be maintained at conventional levels, but the flow rate of cooling air through conduit 104 increased, which can allow the temperature of the working gas to be increased, thereby also improving engine efficiency.

Additionally, the exit position of the cooling air from opening 112, being relatively close to the hub of the engine is advantageous, aerodynamic spoiling losses being generally lower at the hub end of the blade than at its tip.

Figure 4 shows a cross-sectional view of a blade for a turbine stage according to a second embodiment of the invention.

This embodiment is similar to the first embodiment, and corresponding features have the same reference numbers and letters. In this embodiment, however, there are two wdownF! legs to the internal conduit, one extending to an opening 112a at the rear side of the blade, and the other extending to an opening 112b at the front side of the blade.

Again, front and rear annular gaps are formed between platform 109 and stationary portions which in this case are NGV platforms 110, 111. However, the sealing flow B through both the front and the rear gaps is reduced relative to conventional levels as the sealing air flow for the gaps is supplemented by the cooling air exiting openings 112a, 112b.

While the invention has been described in conjunction with the exemplary embodiments described above, many equivalent modifications and variations will be apparent to those skilled in the art when given this disclosure.

Accordingly, the exemplary embodiments of the invention set forth above are considered to be illustrative and not limiting. Various changes to the described embodiments may be made without departing from the spirit and scope of the invention.

Claims

CLAIMS1. A turbine stage of a gas turbine engine, the stage having: an annular passage for conveying a flow of a working gas, a rotor disc, a row of rotor blades mounted at their ends to a rim of the rotor disc, and a corresponding row of nozzle guide vanes; wherein: the blades and vanes extend across the annular passage and, in use, extract power from the working gas, the blades and vanes having internal conduits for carrying a cooling gas flow, the blades have respective platforms adjacent said ends which define a boundary of the annular passage, front and rear edges of the platforms forming corresponding front and rear annular gaps with facing stationary portions of the engine, and the stage is adapted to produce a sealing gas flow which issues from the gaps into the annular passage to prevent the working gas leaking in the opposite direction; wherein at least a portion of the cooling gas flow is reused to form at least a portion of the sealing gas flow.
2. A turbine stage according to claim 1 wherein the internal conduits of the rotor blades route the cooling gas carried therein towards the rim of the rotor disc, the blades having respective openings at said ends through which cooling gas escapes from the internal conduits and forms at least a portion of the sealing gas flow.
3. A turbine stage according to claim 2 wherein the escaped cooling gas flow forms at least a portion of the sealing gas flow through the rear annular gap.
4. A turbine stage according to claim 2 or 3 wherein the escaped cooling gas flow forms at least a portion of the sealing gas flow through the front annular gap.
5. A turbine stage according to any one of the previous claims wherein the stage is adapted so that cooling gas flow which has passed through the internal conduits of the nozzle guide vanes is routed to form at least a portion of the sealing gas flow.
6. A gas turbine engine having the turbine stage of any one of the previous claims.
7. A rotor blade for use in the turbine stage of any one of claims 2 to 4.
8. A nozzle guide vane for use in the turbine stage of claim 5.
9. A turbine stage as any one herein described with reference to and as shown in Figures 3 and 4.
10. A rotor blade as any one herein described with reference to and as shown in Figures 3 and 4.