GB2039627A

GB2039627A - Rotor blade assembly for a gas turbine engine

Info

Publication number: GB2039627A
Application number: GB7910995A
Authority: GB
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 1978-04-04
Filing date: 1979-03-28
Publication date: 1980-08-13

Abstract

The assembly has blades (13) supporting vibration damping devices (20), each comprising a member which extends in a chordwise direction of the blade aerofoil (18) and is attached to the aerofoil at its trailing edge, the leading edge of the member having limited freedom of radial movement sufficient to allow it to engage frictionally with a projection (24) on the aerofoil under the effect of centrifugal force. <IMAGE>

Description

SPECIFICATION A rotor blade assembly for a gas turbine engine This invention relates to a rotor blade assembly for a gas turbine engine.

One of the problems experienced with rotor blades is that they tend to vibrate, which fatigues the blade and can damage other structure of the engine.

Various forms of damping have been used but the vibration still remains a problem particularly in the case of large blades such as fan blades. Previous damping arrangements have used constructions engaged with the root or platform of the blade and have thus been unable to control motion of the aerofoil except in an indirect manner.

The present invention provides a rotor blade assembly having damping means which operates directly on the aerofoil.

According to the present invention a rotor blade assembly for a gas turbine engine has a plurality of blades, each comprising an aerofoil section portion which supports a damping device, the damping device comprising a member which extends in a chordwise direction of the aerofoil and is attached to the aerofoil at one chordwise extent, the other chordwise extent of the member having a limited freedom of radial movement sufficient to allow it to engage frictionally with a projection from the aerofoil under the effect of centrifugal force on the member said engagement providing frictional damping of vibration of the aerofoil.

It is convenient if each damping member spans the gap between two adjacent blades in a rotor stage, thus each damping member may comprise one of a plurality which together make up an annular flow divider or clapper ring.

The invention is most conveniently applied to the fan blades of a gas turbine engine.

The invention will now be particularly described, merely by way of example, with reference to the accompanying drawings in which; Figure 1 is a partly broken away view of a gas turbine engine having fan blades in accordance with the invention, Figure 2 is an enlarged side view of one of the fan blades of Figure 1, Figure 3 is a section on the line 3-3 of Figure 2, Figure 4 is a part section on the line 4-4 of Figure 2, and FigureS is a view on the lineS of Figure 2.

In Figure 1 there is shown a gas turbine engine comprising a fan 10 driven from a core engine 11.

The core engine and the method of driving the fan from the core engine exhaust are conventional. The fan itself comprises a rotor 12 which carries by conventional root engagements an angularly spaced apart row of fan blades 13. The fan blades 13 operate inside a fan cowl 14 and are provided with an annular flow divider ring 15. The ring 15 divides the air flowing through the fan into an inner portion which enters the core engine 11 and an outer portion which passes between the core engine casing and the fan cowl 14to provide propulsive thrust.

Construction of the fan blades 13 is elaborated in Figure 2 where it will be seen that each blade 13 comprises a root 16 which supports, through a shank 17, an aerofoil 18. A platform 19 provides the inner boundary of flow through the fan blade. The flow divider ring 15 can be seen in Figure 2 to comprise a streamlined casing made up of a plurality of partannular casing segments 20. Each of the casing segments 20 is supported from the trailing edge of the blade from a ring 21 which in turn is brazed or otherwise joined to the trailing edge of the aerofoil.

Figure 4 shows that the ring 21 has a series of projections 22 within each of which an indentation 23 engages the trailing edge of the aerofoil 18.

As can best be seen from Figure 5, the remainder of the casing segment 20 is spaced from the outer surface of the aerofoil by a small gap. Therefore so far as the structure described above is concerned, the casing segment 20 would be free to move radially at its forward edge under the influence of centrifugal loads. Projections 24 are therefore provided from either flank of the aerofoil 18 and these projections engage with the upper surface of the lower part of the casing segments 20. The projections 24 therefore prevent any considerable radial movement of the forward portion of the casing segment 20, and more importantly they provide a frictional engagement between the leading edge region of the blade and the adjacent casing segments 20.It will be understood that any vibrational movement of the aerofoil 18 will need to cause relative movement between the frictionally engaging surfaces of the projections 24 and the casing segment 20. This will result in the dissipation of energy and will therefore provide damping of the aerofoil.

Clearly the actual amount of damping provided may be adjusted by varying the size of the projections 24 and the weight of the casing 20. In fact it may be desirable to fill the casing segments 20 with an elastomeric substance and if necessary to provide special bearing surfaces on this material to engage the projections 24.

Clearly the construction described above is particularly advantageous because it utilises the feature which must in any case be present for aerodynamic reasons to provide damping at least part way up the aerofoil. However, it would be possible to apply the invention to blades other than fan blades and in situations where the flow divider is not required.

It should be noted that the construction described shows simply the damping arrangement in accordance with the invention, however it would be quite possible to incorporate more complicated structure into the ring 15. In particular it would be possible to make use of the construction described in our co-pending application 13068/78.

1. A rotor blade assembly for a gas turbine engine having a plurality of rotor blades, each comprising an aerofoil section portion which sup ports a damping device, the damping device com prising a member which extends in a chordwise direction of the aerofoil and is attached to the

**WARNING** end of DESC field may overlap start of CLMS **.

Claims

**WARNING** start of CLMS field may overlap end of DESC **. SPECIFICATION A rotor blade assembly for a gas turbine engine This invention relates to a rotor blade assembly for a gas turbine engine. One of the problems experienced with rotor blades is that they tend to vibrate, which fatigues the blade and can damage other structure of the engine. Various forms of damping have been used but the vibration still remains a problem particularly in the case of large blades such as fan blades. Previous damping arrangements have used constructions engaged with the root or platform of the blade and have thus been unable to control motion of the aerofoil except in an indirect manner. The present invention provides a rotor blade assembly having damping means which operates directly on the aerofoil. According to the present invention a rotor blade assembly for a gas turbine engine has a plurality of blades, each comprising an aerofoil section portion which supports a damping device, the damping device comprising a member which extends in a chordwise direction of the aerofoil and is attached to the aerofoil at one chordwise extent, the other chordwise extent of the member having a limited freedom of radial movement sufficient to allow it to engage frictionally with a projection from the aerofoil under the effect of centrifugal force on the member said engagement providing frictional damping of vibration of the aerofoil. It is convenient if each damping member spans the gap between two adjacent blades in a rotor stage, thus each damping member may comprise one of a plurality which together make up an annular flow divider or clapper ring. The invention is most conveniently applied to the fan blades of a gas turbine engine. The invention will now be particularly described, merely by way of example, with reference to the accompanying drawings in which; Figure 1 is a partly broken away view of a gas turbine engine having fan blades in accordance with the invention, Figure 2 is an enlarged side view of one of the fan blades of Figure 1, Figure 3 is a section on the line 3-3 of Figure 2, Figure 4 is a part section on the line 4-4 of Figure 2, and FigureS is a view on the lineS of Figure 2. In Figure 1 there is shown a gas turbine engine comprising a fan 10 driven from a core engine 11. The core engine and the method of driving the fan from the core engine exhaust are conventional. The fan itself comprises a rotor 12 which carries by conventional root engagements an angularly spaced apart row of fan blades 13. The fan blades 13 operate inside a fan cowl 14 and are provided with an annular flow divider ring 15. The ring 15 divides the air flowing through the fan into an inner portion which enters the core engine 11 and an outer portion which passes between the core engine casing and the fan cowl 14to provide propulsive thrust. Construction of the fan blades 13 is elaborated in Figure 2 where it will be seen that each blade 13 comprises a root 16 which supports, through a shank 17, an aerofoil 18. A platform 19 provides the inner boundary of flow through the fan blade. The flow divider ring 15 can be seen in Figure 2 to comprise a streamlined casing made up of a plurality of partannular casing segments 20. Each of the casing segments 20 is supported from the trailing edge of the blade from a ring 21 which in turn is brazed or otherwise joined to the trailing edge of the aerofoil. Figure 4 shows that the ring 21 has a series of projections 22 within each of which an indentation 23 engages the trailing edge of the aerofoil 18. As can best be seen from Figure 5, the remainder of the casing segment 20 is spaced from the outer surface of the aerofoil by a small gap. Therefore so far as the structure described above is concerned, the casing segment 20 would be free to move radially at its forward edge under the influence of centrifugal loads. Projections 24 are therefore provided from either flank of the aerofoil 18 and these projections engage with the upper surface of the lower part of the casing segments 20. The projections 24 therefore prevent any considerable radial movement of the forward portion of the casing segment 20, and more importantly they provide a frictional engagement between the leading edge region of the blade and the adjacent casing segments 20.It will be understood that any vibrational movement of the aerofoil 18 will need to cause relative movement between the frictionally engaging surfaces of the projections 24 and the casing segment 20. This will result in the dissipation of energy and will therefore provide damping of the aerofoil. Clearly the actual amount of damping provided may be adjusted by varying the size of the projections 24 and the weight of the casing 20. In fact it may be desirable to fill the casing segments 20 with an elastomeric substance and if necessary to provide special bearing surfaces on this material to engage the projections 24. Clearly the construction described above is particularly advantageous because it utilises the feature which must in any case be present for aerodynamic reasons to provide damping at least part way up the aerofoil. However, it would be possible to apply the invention to blades other than fan blades and in situations where the flow divider is not required. It should be noted that the construction described shows simply the damping arrangement in accordance with the invention, however it would be quite possible to incorporate more complicated structure into the ring 15. In particular it would be possible to make use of the construction described in our co-pending application 13068/78. CLAIMS

1. A rotor blade assembly for a gas turbine engine having a plurality of rotor blades, each comprising an aerofoil section portion which sup ports a damping device, the damping device com prising a member which extends in a chordwise direction of the aerofoil and is attached to the aerofoil at one chordwise extent, the other chordwise extent of the member having limited freedom of radial movement sufficient to allow it to engage frictionally with a projection from the aerofoil under the effect of centrifugal force on the member, said engagement providing frictional damping of vibration of the aerofoil.

2. A rotor blade assembly as claimed in claim 1 and in which each damping member is adapted to span the gap between two adjacent said blades.

3. A rotor blade assembly as claimed in claim 2 and in which each damping member frictionally engages with the two adjacent blades with which it is associated.

4. A rotor blade assembly as claimed in claim 3 and in which said damping members together make up a ring.

5. A rotor blade assembly as claimed in claim 4 and in which said ring of damping members is shaped to act as a flow divider.

6. A rotor blade assembly as claimed in any of the preceding claims and in which a supporting ring interconnects the trailing edges of the aerofoil and supports said one chordwise extent of each said damping member from adjacent blades.

7. A rotor blade assembly as claimed in any preceding claim and in which each said damping member comprises a streamlined casing segment.

8. A rotor blade assembly as claimed in claim 7 and in which each said casing segment is filled with elastomeric material.

9. A rotor blade assembly substantially as hereinbefore particularly described with reference to the accompanying drawings.

10. A gas turbine engine having a rotor blade assembly as claimed in any one of the preceding claims.