GB2546481A

GB2546481A - Rotor stage

Info

Publication number: GB2546481A
Application number: GB1600599.3A
Authority: GB
Inventors: Pursell John
Original assignee: Rolls Royce PLC
Current assignee: Rolls Royce PLC
Priority date: 2016-01-13
Filing date: 2016-01-13
Publication date: 2017-07-26
Also published as: GB201600599D0

Abstract

A rotor stage 13 for a gas turbine engine comprises a disc 44, a first set of blades 40 extending radially outwards from the disc and a second set of blades 42 extending radially outwards from the disc. The first set of blades is permanently attached to the disc and the second set of blades is non-permanently attached to the disc. The first and second sets of blades may be arranged in a repeating pattern around the disc, and each blade of the first or second set may be disposed between and adjacent to two blades of the other set. The first set of the blades may be attached by linear friction welding, and the second set of blades may be retained in slots 46, with corresponding root sections 48 of fir-tree, dove-tail or bulb type. The first set of blades may be permanently attached to stubs projecting from the disc. Also claimed are methods of manufacturing, balancing and repairing such a rotor stage, as well as a rotor disc assembly with a first set of permanently attached blades, and slots suitable for receiving a second set of blades.

Description

ROTORSTAGE

Technical Field

The present disclosure concerns a rotor stage for a gas turbine engine, and/or a fan for a gas turbine engine, and/or a bladed disc, and/or a method of manufacturing, balancing and/or repairing a rotor stage.

Background

Gas turbine engines are typically employed to power aircraft. A gas turbine engine comprises various stages of rotor blades which rotate in use. Typically, a gas turbine engine would have at least one compressor rotor stage, and at least one turbine rotor stage. Typically, a gas turbine engine for aerospace purposes includes a fan.

There are a number of ways in which the blades of a rotor stage may be attached to the engine. Generally, the blades attach to a disc that is linked to a shaft. Conventionally, blades are inserted and locked into slots formed in such discs.

There is a drive to make aero engines lighter. Integral bladed disc rotors also referred to as blisks, can reduce the weight of the rotor stage. Such blisks may be, for example, machined from a solid component, or may be manufactured by welding of the blades to the disc. Blisks are therefore increasingly used in modern gas turbine engines, for example as part of the compressor and/or fan.

It is desirable for blisks to be designed with minimal vibration responses from, for example, resonance and flutter, which may be distortion driven.

Summary

According to a first aspect there is provided a rotor stage for a gas turbine engine comprising a disc, a first set of blades extending radially outwards from the disc and a second set of blades extending radially outwards from the disc. The first set of blades may be permanently attached to the disc. The second set of blades may be non-permanently attached to the disc.

In the present application a permanent attachment is an attachment between two components whereby separating the two components may damage either component. Examples of permanent attachment methods include welding (e.g. by linear friction welding), co-moulded (e.g. machined from a single block of material), or being partly or fully integral components.

In the present application a non-permanent attachment is an attachment between components whereby the components can be separated with no, or a reduced, risk of damage to either component.

The first set of blades and the second set of blades may be arranged in a pattern that repeats itself around the entire circumference of the disc.

The disc may have a central axis. The first set of blades and the second set of blades may be arranged circumferentially around the disc. The rotor stage may comprise only blades of the first and second set. There may be an equal number of blades of the first set and the second set. There may be more blades of the first set than the second set. There may be more blades of the second set than the first set. A set of blades may comprise a plurality of blades.

The first set of blades and the second set of blades may be arranged to be axi-symmetric around the disc.

Each blade of the first set of blades may be circumferentially located between, and adjacent to, two blades of the second set of blades.

Each blade of the second set of blades may be circumferentially located between, and adjacent to, two blades of the first set of blades. A blade (e.g. each blade) of the first set of blades may be adjacent to a blade of the second set of blades. A blade (e.g. each blade) of the second set of blades may be adjacent to a blade of the first set of blades.

For example, a blade (e.g. each blade) of the first set of blades may be circumferentially located between, and adjacent to, two blades of the second set of blades.

Blades of the first and second set of blades may be arranged alternately around the entire circumference of the disc.

The blades of the first and second set of blades may be arranged alternately around the entire circumference of the disc such that in a clockwise circumferential direction each blade of the first set of blades is proceeded by a blade of the second set of blades, the blade of the second set of blades itself proceeded by a blade of the first set of blades.

The disc may comprise slots. The slots may comprise flanks. The blades of the second set of blades may comprise blade roots. The blade roots may comprise blade root flanks. Each slot may receive a blade root of a blade of the second set of blades. The slots and the blade roots may be configured such that the blade root is moveable in the slot between a first position and a second position. In a first position the blade root flanks may be in frictional contact with the flank of the slot. In a second position there may be a gap between the blade root flank and the flank of the slot.

Friction between the blade (for example the blade root) and the slot can improve damping performance. A non-permanent attachment may include an interface between the two constrained components where they are in frictional contact. For example a nonpermanent attachment may include inserting and locking a blade into a slot formed in a disc.

The blade root and the slot may be in the first position during operation, for example when the rotor stage is rotating.

The blade roots of the second set of blades may be of dovetail, bulb, or fir tree type.

The first set of blades may be attached to the disc by welding.

The first set of blades may be attached to the disc by linear friction welding.

The disc may comprise stubs that project radially outwards from the disc. Each blade of the first set of blades may be attached (e.g. permanently attached) to one of the stubs.

The stubs may be permanently attached to the rotor disc. For example, the stubs may be integrally formed with the disc. The stubs may be formed by machining and/or forging.

The slots may be defined by two adjacent stubs, one on each side of the slot. The slots may be defined by two adjacent stubs and blade roots of two adjacent blades of the first set.

The rotor stage may comprise annulus fillers. A fan for a gas turbine engine may comprise a rotor stage as described herein.

The rotor stage may be a fan. In this arrangement, the first set of blades and the second set of blades may be fan blades. The rotor disc arrangement allows for a smaller hub-tip ratio that may be of particularly benefit for fan blades. Alternatively, the rotor stage may be a compressor stage or a turbine stage.

The first set of blades may be forged blades. The second set of blades may be forged blades. The first set of blades may be titanium blades. The second set of blades may be titanium blades. The blades may be made from other materials suitable for permanent attachments. All blades within a set may be made of the same material. The blades may be made from the same material as the disc.

The disc may be a solid disc or an annular ring. The disc may be a hub. The inner annulus of the disc may be hobbed. The disc may be linked to a rotatable shaft. The disc may be connected to another rotor stage by a shaft. The disc may be a titanium disc.

According to a second aspect there is provided a method of manufacturing a rotor stage of the first aspect comprising permanently attaching the first set of blades to the disc and non-permanently attaching the second set of blades to the disc.

The first set of blades may be permanently attached to the disc by linear friction welding. The direction of oscillation of the linear friction welding may be in the circumferential direction.

According to a third aspect there is provided a method of manufacturing a rotor stage, the method comprises selectively arranging the second set of blades around the disc such that the variance of the location of the centre of mass, as well as the variance in mass, of the individual blades of the first and second set, is balanced such that the centre of mass of the rotor stage is aligned with a central axis of the disc.

The rotor stage may be the rotor stage of the first aspect.

According to a fourth aspect there is provided method of repairing a rotor stage of the first aspect, wherein at least a blade of the first set of blades is adjacent to a blade of the second set of blades, the method may comprise removing a blade of the second set of blades adjacent to a blade of the first set of blades for providing access to the blade of the first set of blades for repair.

The method of repairing a rotor stage can improve the ease of repair of the blade of the first set of blades where each blade of the first set of blades is located between, and adjacent to, two blades of the second set of blades.

According to a fifth aspect there is provided a fan of a gas turbine engine comprising a fan disc, a first set of fan blades extending radially outwards from the fan disc and a second set of fan blades extending radially outwards from the fan disc. The first set of fan blades may be welded to the disc. The fan disc may comprise slots. Each blade of the second set of fan blades may comprise a blade root that is received at least partially by one of the slots in the disc.

The first set of fan blades and the disc may define a disc-blade assembly. The disc blade assembly may comprise slots, and a blade root of a blade of the second set of blades may be received in a slot of the disc-blade assembly. For example, the disc and the first set of blades may define the slots, or the disc alone may define the slots.

The fan may be a rotor stage of the first aspect.

According to a sixth aspect there is provided a disc-blade assembly (or a bladed disc) for a rotor assembly of a gas turbine engine comprising a disc, a first set of blades extending radially outwards from the disc. The first set of blades may be permanently attached to the disc. The disc-blade assembly may comprise slots for receiving blade roots of a second set of blades.

The disc-blade assembly may form part of the rotor stage of the first aspect or the fan of the fifth aspect.

Brief description of the drawings

Embodiments will now be described by way of example only, with reference to the Figures, in which:

Figure 1 is a sectional side view of a gas turbine engine;

Figure 2 is a view from an axial direction of a rotor stage comprising permanently and non-permanently attached blades;

Figure 3 is a view from an axial direction of the disc and permanently attached blades of Figure 2; and

Figure 4 shows a plan view of a permanently attached blade.

Detailed description

With reference to Figure 1, a gas turbine engine is generally indicated at 10, having a principal and rotational axis 11. The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high-pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and an exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines both the intake 12 and the exhaust nozzle 20.

The gas turbine engine 10 works in the conventional manner so that air entering the intake 12 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place.

The compressed air exhausted from the high-pressure compressor 15 is directed into the combustion equipment 16 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 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high 17, intermediate 18 and low 19 pressure turbines drive respectively the high pressure compressor 15, intermediate pressure compressor 14 and fan 13, each by suitable interconnecting shaft.

Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. By way of example such engines may have an alternative number of interconnecting shafts (e.g. two) and/or an alternative number of compressors and/or turbines. Further the engine may comprise a gearbox provided in the drive train from a turbine to a compressor and/or fan.

Referring to Figure 2, a fan rotor stage is indicated generally at 13. The rotor stage includes a disc 44 which forms an annular ring with a central hole 50. The disc has slots 46 arranged circumferentially around its outer edge. The slots 46 are dovetail in shape. The slots 46 are circumferentially narrower at a radially outer position of the disc 44 compared to a radially inner position of the disc and include angled flanks 46’ that face into the slot.

The rotor stage also has a first set of blades 40 and a second set of blades 42. The first set of blades 40 and the second set of blades 42 are attached to the disc 44 so that they extend outwards in a radial direction and are arranged circumferentially around the disc 44.

The first set of blades 40 are permanently attached to the disc 44 and are referred to from hereon in as permanently attached blades. The permanently attached blades 40 are welded to the disc 44. However, in alternative embodiments the blades may be attached by any suitable permanent attachment method. For example the permanently attached blades may be integrally formed with the disc. In the present example the permanently attached blades 40 form fan blades.

The second set of blades 42 are non-permanently attached to the disc 44 and are referred to from hereon in as non-permanently attached blades. The non-permanently attached blades 42 each have blade roots 48 which are received by the slots 46 in the disc 44 and this attachment constrains the non-permanently attached blades 42 with respect to the disc 44.

The blade roots 48 of the non-permanently attached blades 42 are of dovetail shape in the embodiment shown in Figure 2. The blade roots 48 include upper flanks 48’. The upper flanks 48’ form surfaces whereby a component of the direction normal to the surface is in the radial outer direction, when the blade is in position on the rotor stage. The upper flanks 48’ correspond in shape to the angled flanks 46’ and the upper flanks 48’ can abut against and be in frictional contact with the angled flanks 46’. It will be understood that other blade root shapes will also be possible.

The slots 46 are configured to receive the blade roots 48 of the non-permanently attached blades 42. The dovetail shape of the blade roots 48 and the shape of the slots 46 are configured to prevent the non-permanently attached blades 42 from escaping from the disc 44. When under a suitable centrifugal load, for example when the gas turbine engine 10 is generating thrust for forward momentum, the upper flanks 48’ of the dovetail blade root 48 engage with the slot flanks 46’ and these surfaces are held in frictional contact by the centrifugal forces. The second set of blades may be referred to in the art as slotted blades.

In alternative embodiments other blade roots and slot shapes may be provided. In such embodiments the blade root and disc slot should be shaped so that the blade root can be received and retained in the disc slot. For example the root may be fir tree shaped and the slot may be fir tree shaped, or alternatively the root may be bulb shaped and the slot may be bulb shaped. It will be appreciated to the skilled person that the blade roots and slots can be of any shape that constrains the blade root within a slot whilst providing surfaces (for example the slot flanks 46’ and blade root flanks 48’) that contact and hold the blade in place during operation.

Figure 2 illustrates an arrangement whereby there is an equal number of permanently attached blades 40 and non-permanently attached blades 42 arranged around the disc 44. The blades are arranged in an alternating fashion such that each blade of one type is sandwiched between two blades of the other type e.g. in a circumferential direction each permanently attached blade 40 is proceeded by a non-permanently attached blade 42 and each non-permanently attached blade 42 is proceeded by a permanently attached blade 40.

In Figure 2 it can be seen that the profile of the gas washed part of the blades is the same, and further all blades extend to substantially the same radial position. All of the non-permanently attached blades are the same design and all of the permanently attached blades are the same design. In alternative embodiments the gas washed part of the first set of blades may have a different shape to the gas washed part of the second set of blades and/or the first set of blades may extend to different radial positions than the second set of blades.

It can be seen that because not all blades require slots 46, the disc 44 can be smaller. A smaller disc can reduce weight and/or ease manufacture.

The central hole 50 may be hobbed in a conventional manner.

The blade forgings can create the dovetail “fingers” which form either side of the dovetail slot 46 for the slotted blades and can enhanced high cycle fatigue strength relative to a disc forging.

The frictional contact between the blade root flanks 48’ and the slot flanks 46’ can improve the damping capacity of the rotor disc, compared to a rotor disc having entirely permanently attached blades.

Referring to Figure 3, a rotor stage similar to the rotor stage of Figure 2 is shown but without the non-permanently attached blades 42 or fully formed slots 46. The permanently attached blades 40 include a blade foot 60 that defines a surface of attachment to the disc. The blade foot 60 may define part of, e.g. half of, the dovetail slot for each adjoining non-permanent blade. The disc 44 includes stubs 62. Provision of stubs can improve control the heat affected zone formed when the blade foot is welded to the disc. The radially outer surface of the stub 62 is shaped to be complimentary to a surface of the blade foot 60. The contacting surfaces of the stub 62 and the blade foot 60 may be of the same dimensions, as in the Figure 3 example.

To manufacture the rotor stage shown in Figure 2, a disc forging may be provided and machined to a form a disc that includes stubs 62. The permanently attached blades 40 and/or the non-permanently attached blades 42 may be manufactured by forging. The permanently attached blades 40 and/or the non-permanently attached blades 42 and/or the disc 44 may be made of titanium. Each blade foot 60 of the permanently attached blades 40 is then linear friction welded to a stub 62. Referring to Figures 3 and 4 the direction of oscillation of the linear friction welding is in the circumferential direction as shown by arrow A. The direction of the welding pressure is radially inwards as shown by arrow B.

With the permanently attached blades 40 welded to the stubs 62 the rotor disc will appear as shown in Figure 3. The slots 46 are then machined (or milled) into the disc and permanently attached blade assembly. The non-permanently attached blades 42 are then slotted into the slots 46 to achieve the rotor stage of Figure 2.

The skilled person will appreciate that the blades can be made from other manufacturing processes or materials suitable for the gas turbine environment. The skilled person will appreciate that the blades and disc can be made of materials that are suitable to allow permanent and non-permanent attachments to exist between them.

To balance a rotor stage according to Figure 2 the non-permanently attached blades may be arranged such that the overall centre of mass of the rotor stage is aligned with the rotational axis. Balancing is often required because each blade will have minor variations in the mass of the blade and/or the centre of mass of the blade which can be caused by manufacturing variances.

To repair a non-permanently attached blade 42, the non-permanently attached blade 42 is removed from its slot 46 and the repair performed with the blade separate from the rotor stage.

To repair a permanently attached blade 40 a non-permanently attached blade 42 adjacent to the permanently attached blade 40 can be removed. The permanently attached blade 40 can then be repaired. Purely by way of example, referring to Figure 2, removing non-permanently attached blade 42A allows access to the permanently attached blades 40A and 40B on either side for maintenance and repair. Further, both of the non-permanently attached blades 42A and 42B, that neighbour permanently attached blade 40B, can be removed to allow access to blade 40B for maintenance and repair. As such repair is simplified.

In the described embodiment an equal number of permanently attached and non-permanently attached blades are provided, but in alternative embodiments there may be an unequal number of permanently and non-permanently attached blades. In alternative embodiments, the pattern that the blades form around the disc may be different to that shown in Figure 2. The pattern may include, for example, two blades of a certain type adjacent to each other.

In alternative embodiments there may be annulus fillers between the blades.

It will be understood that the rotor stage disclosed can also be a compressor stage or a turbine stage.

The rotor stage described herein can improve manufacturability, and/or lightweight construction, and/or damping performance, and / or repairability. The rotor stage described herein may provide under rim damping. Friction between the blade (for example the blade root) and the slot can improve damping performance.

It will be understood that the invention is not limited to the embodiments above-described and various modifications and improvements can be made without departing from the concepts described herein. Except where mutually exclusive, any of the features may be employed separately or in combination with any other features and the disclosure extends to and includes all combinations and subcombinations of one or more features described herein.

Claims

1. A rotor stage for a gas turbine engine comprising: a disc; a first set of blades extending radially outwards from the disc; and a second set of blades extending radially outwards from the disc; wherein the first set of blades is permanently attached to the disc; and the second set of blades is non-permanently attached to the disc.

2. A rotor stage according to claim 1, wherein the first set of blades and the second set of blades are arranged in a pattern that repeats itself around the entire circumference of the disc.

3. A rotor stage according to claim 1 or claim 2, wherein each blade of the first set of blades is circumferentially located between, and adjacent to, two blades of the second set of blades.

4. A rotor stage according to any one of the previous claims, wherein each blade of the second set of blades is circumferentially located between, and adjacent to, two blades of the first set of blades.

5. A rotor stage according to any one of the previous claims, wherein blades of the first and second set of blades are arranged alternately around the entire circumference of the disc.

6. A rotor stage according to any one of the previous claims, wherein the disc comprises slots; and the slots comprise flanks; and the blades of the second set of blades comprise blade roots; and the blade roots comprise blade root flanks; wherein each slot receives a blade root of a blade of the second set of blades; and wherein the slots and the blade roots are configured such that the blade root is moveable in the slot between a first position and a second position; in a first position the blade root flanks are in frictional contact with the flank of the slot; and in a second position there is a gap between the blade root flank and the flank of the slot.

7. A rotor stage according to claim 6, wherein the blade roots of the second set of blades are of dovetail, bulb, or fir tree type.

8. A rotor stage according to any one of the previous claims, wherein the first set of blades are attached to the disc by welding.

9. A rotor stage according to any one of the previous claims, wherein the first set of blades is attached to the disc by linear friction welding.

10. A rotor stage according to any one of the previous claims, wherein the disc comprises stubs that project radially outwards from the disc; wherein each blade of the first set of blades is attached to one of the stubs.

11 .A fan for a gas turbine engine comprising the rotor stage according to any one of the previous claims.

12. A method of manufacturing a rotor stage according to any one of the previous claims, the method comprising: permanently attaching the first set of blades to the disc; and non-permanently attaching the second set of blades to the disc.

13. A method of manufacturing a rotor stage according to claim 12, wherein the first set of blades are permanently attached to the disc by linear friction welding; and the direction of oscillation of the linear friction welding is in the circumferential direction.

14. A method of balancing a rotor stage according to any one of claims 1 to 10; the method comprising: selectively arranging the second set of blades around the disc such that the variance of the location of the centre of mass, as well as the variance in mass, of the individual blades of the first and second set, is balanced such that the centre of mass of the rotor stage is aligned with a central axis of the disc.

15. A method of repairing a rotor stage according to any one of claims 1 to 10, wherein at least a blade of the first set of blades is adjacent to a blade of the second set of blades, the method comprising: removing a blade of the second set of blades adjacent to a blade of the first set of blades for providing access to the blade of the first set of blades for repair.

16. A fan of a gas turbine engine comprising: a fan disc; a first set of fan blades extending radially outwards from the fan disc; and a second set of fan blades extending radially outwards from the fan disc; wherein the first set of fan blades is welded to the disc to form a disc-blade assembly; and the disc-blade assembly comprises slots; wherein each blade of the second set of fan blades comprises a blade root that is received by one of the slots in the disc-blade assembly.

17. A disc-blade assembly for a rotor assembly of a gas turbine engine comprising: a disc; a first set of blades extending radially outwards from the disc; and slots for receiving blade roots of a second set of blades; wherein the first set of blades is permanently attached to the disc.

18. A rotor stage, gas turbine engine and/or method substantially as herein described and/or as shown in the accompanying drawings.