GB2106996A

GB2106996A - Cooled rotor aerofoil blade for a gas turbine engine

Info

Publication number: GB2106996A
Application number: GB08129466A
Authority: GB
Inventors: Michael John Buller
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1981-09-30
Filing date: 1981-09-30
Publication date: 1983-04-20
Also published as: FR2513695A1; JPS5867904A; DE3234906A1

Abstract

A hollow cooled rotor aerofoil blade (10) for a gas turbine engine is provided with an apertured plate (17) which divides the interior of the blade into two portions (15, 16). A major portion of the apertured plate (17) is free and inclined with respect to the blade (10) so that upon rotation of the blade (10) in a gas turbine engine, the apertured plate (17) is centrifugally urged into sealing engagement with inwardly extending ribs (21) provided on the interior surface of the blade (10). One of the portions (16) of the blade interior is supplied with cooling air which then passes through the apertures (22) in the apertured plate (17) into the other portion (15) to provide impingement cooling of the region of the blade (10) around its leading edge (23). The construction also produced a damping effect as well as allowing for temperature differentials. <IMAGE>

Description

SPECIFICATION Cooled rotor aerofoil blade for a gas turbine engine This invention relates to a cooled rotor aerofoil blade for a gas turbine engine.

Gas turbine engine rotor aerofoil blades are frequently required to operate at temperatures which are so high that some form of blade cooling is necessary. One particularly convenient method of blade cooling entails passing cooling air through passages within the blade. The cooling air may be derived from any convenient source but is usually tapped from the higher pressure stages of the compressor of the gas turbine engine. The cooling air may be directed through the blade along any one of a large number of different paths depending upon the desired cooling pattern. If certain areas of the aerofoil portion of the blade are prone to localised overheating, those areas may be cooled by what is commonly referred to as "impingement cooling".The technique of impingement cooling involves directing jets of cooling air at the inner surface of a hollow blade in those areas which are required to be cooled.

There are several ways in which such jets of cooling air may be derived. One way is to form a hollow blade by, for instance, casting, which is provided with an integral apertured wall separating the area to be cooled from a passage carrying cooling air. Cooling air from the passage passes through the apertured wall to provide cooling air jets which impinge upon the area to cooled. The disadvantage of such an arrangement is that the integral apertured wall tends to be overcooled with respect to the rest of the blade and is consequently prone to cracking as a result of thermal stress.

An alternative way in which the cooling air jets may be derived is by the provision of an insert within a hollow blade. The insert is in the form of a tube which is supplied with cooling air and has a number of small holes in it to direct the cooling air on to the areas of the blade which are to be cooled. Inserts are not particularly attractive in view of the weight increase they bring to the blade, the difficulty in effectively anchoring the tube within the blade and possible compromises in the aerodynamic profile of the aerofoil portion of the blade in order to accommodate the insert. It is an object of the present invention to provide an impingement cooled rotor aerofoil blade for a gas turbine engine which provides effective impingement cooling and substantially avoids the aforementioned difficulties.

According to the present invention, a hollow cooled rotor aerofoil blade for a gas turbine engine is provided with an apertured plate adapted to divide the interior of said blade into two portions, at least a major portion of said apertured plate being free and inclined with respect to said blade so that in operation, said at least a major portion of said apertured plate is centrifugally urged into sealing engagement with inwardly extending ribs provided on the interior surface of said blade, one of said portions of said hollow blade interior being adapted to be supplied with cooling air so that said cooling air passes through the apertures in said apertured plate to provide impingement cooling of at least part of the interior surface of the other portion of said blade interior.

The invention will now be described by way of example with reference to the accompanying drawings in which: Fig. 1 is a side view of a hollow cooled rotor aerofoil blade in accordance with the present invention viewed on the section line A-A of Fig. 2.

Fig. 2 is a view on the section line B-B of Fig. 1.

With reference to Fig. 1, a hollow cooled rotor aerofoil blade generally indicated at 10 comprises a root portion 11 which is of conventional form for the attachment of the blade 10 to the periphery of a rotor disc (not shown), an aerofoil portion 1 2 and a platform 1 3 which interconnects the root and aerofoil portions 11 and 12 respectively.

The interior 14 of the hollow blade 10 is divided into two portions 1 5 and 16 by an apertured plate 1 7. The plate 1 7 is attached to the aerofoil portion 12 only at the blade tip 1 8. More specifically the radially outer portion 1 9 of the apertured plate 1 7 is bent over to engage a location feature 20 provided at the radially outer end of the blade interior portion 1 5. The radially outer portion 19 of the apertured plate 1 7 is welded (although it could be brazed) to the location feature 20 to leave the remainder of the apertured plate 1 7 free with respect to the blade 10.Moreover, the free portion of the apertured plate 1 7 is inclined with respect to the blade 10 so that upon rotation of the blade 10, the free portion of the apertured plate 1 7 will deflect under centrifugal loading about the location feature 20.

Deflection of the apertured plate 1 7 is limited by inwardly extending ribs 21 provided on the interior surface of the blade 1 0. The ribs 21 provided on the interior surfaces of both of the flanks of the blade 10 and also adjacent the radially inner end of the apertured plate 17 so that the ribs 21 together provide a seal around the periphery of the apertured plate 1 7. Thus rotation of the blade 10 urges the free portion of the apertured plate 1 7 into sealing engagement with the ribs 21, thereby ensuring that the apertures 22 provide the only communication between the blade interior portions 1 5 and 1 6.

The blade interior portion 1 5 is adjacent the leading edge 23 of the aerofoil portion 12 and constitutes a minor portion of the total interior volume of the blade 10. The blade interior portion 1 6 is constituted by the remainder of the interior of the aerofoil portion 12 and additionally includes the interior of the root porfion 11.

In operation cooling air is supplied to the blade interior portion 1 6 from a convenient source such as the compressor of the gas turbine engine in which the blade 10 is mounted. The cooling air enters the blade interior portion 16 through the root 11 and provides radial flow convection cooling of those portions of the aerofoil 12 which serve to define the blade interior portion 1 6. Some of the cooling air is exhausted from the blade interior portion 1 6 through an aperture 24 provided in a cap 25 which encloses the blade interior portion 16 at the blade tip 18.The remainder of the cooling air, however, passes through the apertures 22 provided in the plate 1 7 to form jets which impinge upon the interior surface of the blade leading edge 23, thereby providing impingement cooling of the leading edge 23. The cooling air then passes through the blade interior portion 1 5 providing a certain degree of convection cooling before being exhausted from the blade interior portion 1 5 through an aperture 26 provided in the radially outer, bent-over portion 1 9 of the apertured plate 17.

Although the cooling air has been described as being exhausted from the blade interior portions 15 and 16 through apertures 24 and 26 at the blade tip 18, apertures could in fact be provided on the flanks of the aerofoil portion 12 so that cooling air exhausted through them would provide film cooling of the outer surface of the aerofoil portion 12.

Since the apertured plate 1 7 is centrifugally urged into sealing engagement with the ribs 21, leakage of cooling air between the apertured plate 17 and the ribs 21 is minimised. The present invention thereby ensures that an effective seal is achieved between the apertured plate 1 7 and the ribs 21. Moreover, since the major portion of the apertured plate 1 7 is not fixedly attached to the blade 10, very large temperature difference between the blade 10 and the apertured plate 17 may be tolerated without the occurrence of thermaliy induced stress.

An additional advantage of the present invention is that frictional engagement between the apertured plate 1 7 and the ribs 21 ensures that the plate 17 provides a certain degree of damping of vibration within the blade 1 0.

It will be appreciated that the present invention is not specifically limited to the impingement cooling of the leading edge of a rotary aerofoil blade. Indeed the apertured plate 1 7 could be located in other positions within the blade 10 so as to provide cooling of areas on, for instance, the blade flanks. The apertured plate 17 would, of course, have to be inclined with respect to the blade 10 in such a manner as to ensure that it is centrifugally urged into sealing engagement with appropriately located ribs on the internal surface of the blade 10.

It will also be appreciated that the apertured plate 1 7 need not necessarily be fixedly attached at the tip 1 8 region of the blade 10 but could for instance be fixedly attached to the blade 10 in the region of its platform 1 3. Moreover, the apertured plate need not necessarily be fixedly attached to the blade 10 by brazing or welding but could alternatively be attached by some suitable form of mechanical connection.

Claims

1. A hollow cooled rotor aerofoil blade for a gas turbine engine provided with an apertured plate adapted to divide the interior of said blade into two portions, at least a major portion of said apertured plate being free and inclined with respect to said blade so that in operation said at least a major portion of said apertured plate is centrifugally urged into sealing engagement with inwardly extending ribs provided on the interior surface of said blade, one of said portions of said hollow blade interior being adapted to be supplied with cooling air so that said cooling air passes through the apertures in said apertured plate to provide impingement cooling at least part of the interior surface of the other portion of said blade interior.

2. A hollow aerofoil blade as claimed in claim 1 wherein a minor portion of said apertured plate is fixedly attached to said blade in the region of the radially outer tip thereof.

3. A hollow aerofoil blade as claimed in claim 2 wherein said minor portion of said apertured plate is brazed or welded to said blade in the region of the radially outer tip thereof.

4. A hollow aerofoil blade as claimed in any one preceding claim wherein said part of the interior surface of said blade interior which is impingement cooled is adjacent the leading edge of said blade.

5. A hollow aerofoil blade substantially as hereinbefore described with reference to and as shown in the accompanying drawings.