GB2250548A

GB2250548A - Cooled turbine aerofoil blade

Info

Publication number: GB2250548A
Application number: GB9026548A
Authority: GB
Inventors: John Leslie Winter
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1990-12-06
Filing date: 1990-12-06
Publication date: 1992-06-10
Also published as: GB9026548D0

Abstract

A cooled turbine aerofoil blade (20) is provided with a cooling air passage (26) which is of generally labyrinthine form. Cooling air is supplied to the cooling air passage (26) through two inlets (28), (29), one positioned downstream of the other. The second cooling air inlet (29) boosts the air pressure within the cooling air passage (26) so that cooling effectiveness is maintained along the full length of the cooling air passage (26). Cooling air is also supplied to trailing edge passage 37 via an inlet 35 and film cooling holes are provided. <IMAGE>

Description

COOLED TURBINE AEROFOIL BLADE This invention relates to a cooled turbine aerofoil blade suitable for use in a gas turbine engine.

The high temperatures which operationally exist in the turbine of a gas turbine engine necessitate the provision of some form of cooling of the aerofoil blades within the turbine. Typically the aerofoil blades are provided with internal passages through which cooling air is directed.

The internal passages are fed with cooling air through inlets provided in the radially inward region of the blades.

The passages then direct the cooling air along sinuous paths within the aerofoil portions of the blades to provide cooling thereof. Some of the cooling air is exhausted from the aerofoil portions through holes provided in the aerofoil surfaces to provide additional cooling of those surfaces.

The remainder is typically exhausted through holes provided in the aerofoil tip and trailing edge regions.

It has been found with cooled aerofoil blades of this type that the path length of the cooling air through the blade tends to be long. This results in the air pressure within the passages falling to unacceptably low levels at locations remote from the position of cooling air entry.

It is an object of the present invention to provide a cooled turbine aerofoil blade in which the undesirable effects of such falls in air pressure within the passages are reduced.

According to the present invention, a cooled turbine aerofoil blade comprises a root portion for the attachment of said blade to a rotor disc, an aerofoil portion, and a shank portion interconnecting said root and aerofoil portions, said aerofoil portion containing a cooling fluid passage of labyrinthine form, said labyrinthine form passage having at least two inlets for the flow of cooling fluid thereinto, one of said inlets being located at a position upstream, with respect to the flow of cooling fluid through said labyrinthine form passage, of the other or others of said inlets.

The invention will now be described, by way of example, with reference to the accompanying drawings in which: Figure 1 is a sectioned side view of the upper half of a ducted fan gas turbine engine provided with a cooled turbine aerofoil blade in accordance with the present invention.

Figure 2 is partially broken away perspective view of a cooled turbine aerofoil blade in accordance with the present invention.

Referring to Figure 1, a ducted fan gas turbine engine generally indicated at 10 comprises, in axial flow series, an air intake 11, a ducted fan 12, an intermediate pressure compressor 13, a high pressure compressor 14, combustion equipment 15, a high pressure turbine 16, an intermediate pressure turbine 17, a low pressure turbine 18 and an exhaust nozzle 19.

Air taken in through the intake 11 is accelerated by the fan 12. Part of that air flow is exhausted from the engine 10 to provide propulsive thrust while the remainder is directed into the intermediate pressure compressor 13.

There the air is compressed before being directed into the high pressure compressor 14 where further compression takes place. The compressed air flows into the combustion equipment 15 where it is mixed with fuel and the mixture combusted. The resultant high temperature combustion products then expand through and thereby drive the high, intermediate and low pressure turbines 16,17 and 18 before being exhausted through the exhaust nozzle 19 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 16,17 and 18 respectively drive the high and intermediate pressure compressors 14 and 13 and the fan 12.

The present invention is particularly concerned with the annular array of aerofoil blades of the rotor stage of the high pressure turbine 16.

Each aerofoil rotor blade of the high pressure turbine 16, one of which 20 can be seen more clearly in Figure 2, is of similar generally conventional external configuration.

It comprises a root 21, a shank 22, a platform 23, an aerofoil portion 24 and a shrouded tip 25. The shrouded tip 25 and platform 23 cooperate with the shrouded tips and platforms of circumferentially adjacent aerofoil blades 20 to define a portion of the gas passage which contains the aerofoil portions 24. The root portion 21 is interconnected with the remainder of the blade 20 by means of the shank 22.

The root portion 21 is of the conventional "fir-tree" configuration to facilitate its attachment to the rim of a rotor disc (not shown).

The aerofoil portion 24 contains a primary cooling air passage 26 which is of generally labyrinthine form. Thus the primary cooling air passage 26 makes five longitudinal passes of the aerofoil portion 24; the passes being generally parallel with each other. High pressure cooling air derived from the high pressure compressor 14 is directed to the primary cooling air passage 26 through a feed passage 27 which extends through the blade root 21 and shank 22.

Some cooling air is directed into the primary cooling air passage 26 through a first inlet 28 provided at its most upstream point. Further cooling air is directed into the primary cooling air passage from the feed passage 27 through a second inlet 29 at a location downstream of the first inlet 28. Thus whereas the first inlet 28 directs cooling air into the first pass of the primary cooling air passage 26, the second inlet 29 directs cooling air into the third pass of the cooling air passage 26.

The major outlet 30 of the primary cooling air passage 26 is located in the shrouded tip 25. However a large number of subsidiary outlets 31 are provided in the final two passes of the primary cooling air passage 26. The subsidiary outlets 31 exhaust some of the cooling air from the primary cooling air passage 26 on to the external surface of the aerofoil portion 24, thereby providing film cooling of that surface. The cooling air directed into the primary cooling air passage 26 therefore provides cooling of the aerofoil portion 24 by both convection cooling as it passes through the passage 26 and film cooling as it flows over the external surface of the aerofoil portion 24.

Clearly since the primary cooling air passage 26 is long, there is an inevitable drop in air pressure along the passage length. However, the provision of the second cooling air inlet 29 ensures that the pressure of the air within the primary cooling air passage 26 is boosted part way along its total extent. Consequently sufficient air pressure is available in the downstream regions of the primary cooling air passage 26 for the air to continue to provide effective aerofoil cooling.

The cooling air feed passage 27 additionally directs high pressure cooling air into a longitudinally extending passage 32 provided in the leading edge region 33 of the blade aerofoil portion 24. That air is exhausted through film cooling outlets 34 provided along the leading edge region 33 and also through an outlet 35 provided in the shrouded tip 25. Consequently, the leading edge region 33 is also provided with both film and convection cooling.

Cooling air also derived from the high pressure compressor 14 but at a lower pressure than that directed into the feed passage 27 is directed through an inlet 35 provided adjacent the platform 23 underside into the trailing edge region 36 of the aerofoil portion 24. The cooling air is caused to flow through a longitudinally extending passage 37 in the trailing edge region 36. That cooling air is exhausted from the passage 37 through a plurality of holes 38 provided in the trailing edge region 36 and through an outlet 39 provided in the tip shroud 25.

It will be seen therefore that the aerofoil portion 24 of the blade 20 is cooled by high pressure air flowing through the primary cooling air passage 26 and the leading edge passage 32, and lower pressure air flowing through the trailing edge passage 37.

Although the present invention has been described with reference to a turbine blade 20 having a labyrinthine cooling passage 26 which is provided with a single inlet 29 for boosting air pressure within the passage 26, it will be appreciated that additional pressure boost inlets could be provided if necessary.

Moreover although the present invention has been described with reference to a turbine rotor blade, it is also applicable to turbine stator vanes. According references to blades should be construed as extending to stator vanes.

Claims

Claims:

1. A cooled turbine aerofoil blade comprising a root portion for the attachment of said blade to a rotor disc, an aerofoil portion, and a shank portion interconnecting said root and aerofoil portions, said aerofoil portion containing a cooling fluid passage of labyrinthine form, said labyrinthine form passage having at least two inlets for the flow of cooling fluid thereinto, one of said inlets being located at a position upstream, with respect to the flow of cooling fluid through said labyrinthine passage of the other or others of said inlets.

2. A cooled turbine aerofoil blade as claimed in claim 1 wherein said cooling fluid passage is constituted by a plurality of generally longitudinally extending portions, which are generally parallel with each other.

3. A cooled turbine aerofoil blade as claimed in claim 1 or claim 2 wherein said cooling fluid passage is provided with a plurality of holes which facilitate the flow of at least some cooling fluid from said cooling fluid passage to the exterior surface of said aerofoil portion to provide film cooling thereof.

4. A cooled turbine aerofoil blade as claimed in any one preceding claim wherein said at least two inlets for the flow of cooling fluid into said cooling fluid passage are supplied with cooling fluid from a common cooling fluid feed passage.

5. A cooled turbine aerofoil blade as claimed in claim 4 wherein said common cooling fluid feed passage is located in the root and shank of said blade.

6. A cooled turbine aerofoil blade as claimed in claim 5 wherein said common cooling fluid supply passage supplies cooling fluid to at least one additional cooling fluid passage in said aerofoil portion.

7. A cooled turbine aerofoil blade as claimed in claim 6 wherein said additional cooling fluid passage is located in the leading edge region of said blade.

8. A cooled turbine aerofoil blade as claimed in any one preceding claim wherein at least one further cooling fluid passage is provided in said aerofoil portion, said further cooling fluid passage being supplied with cooling fluid from a source other than said common cooling fluid supply passage.

9. A cooled turbine aerofoil blade as claimed in claim 8 wherein said at least one further cooling fluid passage is or are located in the trailing edge region of said blade.

10. A cooled turbine blade as claimed in claim 2 wherein said cooling fluid passage of labyrinthine form is provided with five of said generally longitudinally extending portions.

11. A cooled turbine blade as claimed in claim 10 wherein one of said cooling fluid inlets is positioned to direct cooling fluid into the first of said generally longitudinally extending passage portions and a further one of said cooling fluid inlets is positioned to direct cooling fluid into the third of said generally longitudinally extending passage portions.

12. A cooled turbine blade substantially as hereinbefore described with reference to and as shown in the accompanying drawings.