GB2607886A

GB2607886A - Rotor assembly and method of assembling a rotor assembly for a gas turbine engine

Info

Publication number: GB2607886A
Application number: GB2108382.9A
Authority: GB
Inventors: Mathew Walker Paul
Original assignee: Siemens Energy Global GmbH and Co KG
Current assignee: Siemens Energy Global GmbH and Co KG
Priority date: 2021-06-11
Filing date: 2021-06-11
Publication date: 2022-12-21
Also published as: WO2022258257A1; GB202108382D0

Abstract

A rotor assembly 70, for a gas turbine engine (10, fig 1), comprises a disc 36 and at least one blade 38. The disc has at least one slot 72 with at least one pair of bearing surfaces 76, 78, 80, 82. The blade has a root portion 84 having a root fixture 86, with at least one pair of bearing surfaces 98, 100, 102, 104, and a radially outwardly platform 60, with a radially inner surface 64. A sleeve 110 is located between the slot and the root fixture, in contact with the slot’s bearing surfaces and the root fixture’s bearing surfaces. The sleeve extends and is in contact with the radially inner surface of the platform. The sleeve and the blade may be made from different superalloys each having a different composition of chromium (Cr). The sleeve is formed monolithically by a casting process or an additive manufacture process.

Description

ROTOR ASSEMBLY AND METHOD OF ASSEMBLING A ROTOR ASSEMBLY FOR A GAS TURBINE ENGINE

FIELD OF INVENTION

The present invention relates to a blade assembly for a rotor assembly for a gas turbine engine.

BACKGROUND OF INVENTION

In a gas turbine engine, the turbine is subject to very high temperatures of working gas from the combustor. The turbine, particularly the first two rotor blade stages are subject to environments that can degrade the components by means of corrosion attack. The corrosion attack preferentially attacks the gamma prime, the primary strengthening phase in nickel-based superalloys, in the substrate resulting in a reduced service-life because of cracking and eventually resulting in a blade failure in the worst cases. The corrosion is driven by gas-born salt deposits and Sulphur, which are well known to accelerate corrosion attack of nickel-based alloys such as those used in turbine blades and vanes.

It is well known to apply corrosion-resistant coatings to the surfaces of turbine blades. However, application of coatings can alter the mechanical properties of turbine blades and particularly in the root fixture where very high stresses are incurred during use.

SUMMARY OF INVENTION

To address the problems of corrosion-attack and blade degradation there is provided a rotor assembly for a gas turbine engine, the rotor assembly comprises a rotational axis, a disc and at least one blade. The disc comprises at least one slot. The slot has at least one pair of bearing surfaces. The blade comprises a root portion having a root fixture and radially outwardly a platform. The platform has a radially inner surface. The root fixture comprises at least one pair of bearing surfaces. A sleeve is located at least partly between the slot and the root fixture. The sleeve is in contact with the slot's bearing surfaces and the root fixture's bearing surfaces. The sleeve extends and is in contact with at least a part of the radially inner surface. The sleeve prevents deposits of harmful substances on surfaces of the blade which might otherwise cause corrosion of the blade below the platform. The sleeve enables the blade to be formed having a single crystal structure and with a relatively low chromium content. Such a single crystal blade with low Cr content is particularly robust and particularly resistant to creep. Low Cr content superalloys are susceptible to corrosion. Parts of the blade exposed to hot working gases are less susceptible to corrosive particle deposits.

The root portion may comprise a neck, the neck is between the root fixture and the platform. The sleeve may be in contact with the neck.

The sleeve may be formed of a superalloy comprising a first composition having a Cr content % by weight. The blade may be formed of a superalloy comprising a second composition having a Cr content % by weight. The second composition may have a Cr content % by weight greater than the first composition having a Cr content % by weight.

The second composition may have a Cr content % by weight greater than 3% Cr the first composition having a Cr content % by weight.

The sleeve may be formed from the first composition having a Cr content at least 12% by weight. The sleeve may be formed from the first composition having a Cr content less than 25% by weight. The sleeve may be formed from the first composition having a Cr content approximately 22% by weight.

The blade may be formed from the second composition having a Cr content between 3% and 18% by weight. The blade may be formed from the second composition having a Cr content less than 12%.

Although the % ranges of possible Cr content in the blade and sleeve overlap, for any one blade and sleeve assembly the Cr content of the sleeve is always greater than that of the blade.

A line may have a constant radius from the rotational axis and may pass through the radially outward edge of the bearing surfaces of the slot. The sleeve may be in contact with substantially all the surface of the slot radially inwardly of the line.

A line may have a constant radius from the rotational axis and may pass through the radially outward edge of the bearing surfaces of the root portion. The sleeve may be in contact with substantially all the surface of the root portion radially inwardly of the line.

The root portion may comprise a radially inner surface and a cooling passage through the radially inner surface. The sleeve may comprise a cooling air inlet that is aligned with cooling passage to allow cooling air to pass into the root portion.

The sleeve may comprise an end plate. Preferably the end plate is at a leading end of the sleeve. The end plate is useful to locate the sleeve to the blade and to protect the end of the blade's root fixture against corrosive deposits.

The method may comprise inserting the sleeve into the slot and inserting the root fixture into the sleeve.

The method may comprise peening the sleeve against the bearing surfaces of the slot prior to inserting the root fixture into the sleeve.

The method may comprise inserting the root fixture into the sleeve and inserting the root fixture and the sleeve into the slot.

The method may comprise peening the sleeve against the bearing surfaces of the root fixture prior to inserting the root fixture and the sleeve into the slot.

The sleeve may be formed monolithically by a casting process or an additive manufacturing process.

BRIEF DESCRIPTION OF THE DRAWINGS

The above-mentioned attributes and other features and advantages of this invention and the manner of attaining them will become more apparent and the invention itself will be better understood by reference to the following description of embodiments of the invention taken in conjunction with the accompanying drawings, wherein FIG. 1 shows part of a turbine engine in a sectional view and in which a rotor assembly in accordance with the present invention is incorporated, FIG. 2 shows a first embodiment of a radial section through a rotor assembly in accordance with the present invention, FIG. 3 shows a second embodiment of a radial section through a rotor assembly in accordance with the present invention, FIG. 4 is a perspective view of a sleeve which is part of the rotor assembly and in accordance with the present invention, FIG. 5 is an enlarged view on area A shown in FIG. 3.

DETAILED DESCRIPTION OF INVENTION

FIG. 1 shows an example of a gas turbine engine 10 in a sectional view. The gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20. The gas turbine engine 10 further comprises a shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10. The shaft 22 drivingly connects the turbine section 18 to the compressor section 14.

In operation of the gas turbine engine 10, air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16. The burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28. The combustion chambers 28 and the burners 30 are located inside the burner plenum 26. The compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.

This exemplary gas turbine engine 10 has a cannular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment. An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.

The turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs. In addition, guiding vanes 40, which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38. The turbine section 18 comprises a turbine casing 51 that surrounds the turbine blades 38 and defines a radially outer surface 55 of a working gas path 57. A radially inner surface 59 further defines the working gas path 57. Parts of the radially inner surface 59 and the radially outer surface 55 are defined by platforms of the blades and stator vanes as is well known. A tip gap is defined between the tip of the blades 38 and the casing 51 or radially outer surface 55.

The combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22. The guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.

The turbine section 18 drives the compressor section 14. The compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.

The rotor blade stages 48 comprise a rotor disc supporting an annular array of blades. The compressor section 14 also comprises a compressor casing 50 that surrounds the rotor stages and supports the vane stages 46. The guide vane stages include an annular array of radially extending vanes that are mounted to the compressor casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point. Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.

The compressor casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14. A radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor and which is partly defined by the annular array of blades 48.

The present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine. However, it should be appreciated that the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.

The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated. The terms forward and rearward refer to the general flow of gas through the engine. The terms axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.

Figures 2 and 3 are sections through part of a rotor assembly 70 and are viewed in the axial direction. Figures 2 and 3 show first and second embodiments of the rotor assembly 70 respectively. The rotor assembly 70 comprises a disc 36 and when fully assembled an annular array of radially extending blades 38, only one of which is shown. The disc 36 comprises an annular array of disc slots 72, only one of which is shown, evenly spaced around the radially outer periphery 74 of the disc 36.

Each blade 38 comprises a root portion 84 having a root fixture 86 and a neck 88.

The neck 88 may be optional on other blades. The root fixture 86, and where applicable the neck 88, blends into a platform 60 which has a radially inner surface 64 and a radially outer surface 62 or gas washed surface 62. Extending, in a radially outward direction, from the platform 60 is an aerofoil 66 comprising a leading edge and a tailing edge. A generally concave pressure surface and a generally convex suction surface each extend between the leading and trailing edges in conventional manner. The aerofoil's 66 pressure and suction surfaces also extend in a radial direction from the platform 60 to a tip of the aerofoil 66 as is well known. The blade 80 shown here is an unshrouded blade, but in other examples the blade 80 could be a shrouded blade or a blade having one or more winglets as known in the art. A tip gap is formed between the tip of the blade and the casing 51.

In the first embodiment shown in Figure 2, the root fixture 86 comprises a fir-tree root configuration and the slot 72 comprises a generally corresponding shape as is well known. In this fir-tree root configuration each disc slot 72 has two (or more) pairs of bearing surfaces 76, 78 and 80, 82 where one bearing surface 76, 80 opposes the other bearing surface 78, 82 of its pair respectively. The root fixture 86 comprises two pairs of lobes 90, 92 and 94, 96 respectively. Each lobe 90, 92, 94, 96 defines a bearing surface 98, 100, 102, 104 respectively. In use, loads from the blade 38 are transmitted through corresponding bearing surfaces, that is the blade's bearing surfaces 98, 100, 102, 104 and the slot's bearing surfaces 76, 78, 80, 82 respectively.

In the second embodiment, shown in Figure 3, the root fixture 86 is a single lobe dovetail design and the slot 72 has a generally corresponding shape as is well known. The slot 72 has a pair of bearing surfaces 76, 78 where one bearing surface 76 opposes the other bearing surface 78. The root fixture 86 comprises a pair of lobes 90, 92 respectively. Each lobe 90, 92 defines a bearing surface 98, 100 respectively. In use, loads from the blade 38 are transmitted through corresponding bearing surfaces, that is the blade's bearing surfaces 98, 100 and the slot's bearing surfaces 76, 78 respectively.

In both embodiments shown in Figures 2 and 3, the rotor assembly 70 further comprises a sleeve 110 disposed at least partly between the root fixture 86 and the slot 72. In particular, the sleeve 110 is located between the blade's bearing surfaces 98, 100, 102, 104 and the slot's bearing surfaces 76, 78, 80, 82 and as such transmits loads, in use, from the blade 38 to the disc 72 via the respective bearing surfaces.

In the first embodiment shown in Figure 2, the sleeve 110 is assembled to the slot 72 first. The sleeve 110 has an external surface 114 that is in contact with the surface 73 of the slot 72 radially inwardly of the radially outermost part of the bearing surfaces 76, 78 as shown by line 112. Line 112 is defined by a constant radius from the rotational axis 20 and intersects the radially outermost part of the bearing surfaces 76, 78 of the slot 72. Dashed line R denotes a radial line from the engine rotational axis 20. When the blade 38 is assembled into the sleeve 110, an internal surface 116 of the sleeve 110 is in contact with the surface of the neck 88 and the radially inner surface 64 of the platform 64 and, radially inwardly of the line 112, the sleeve 110 only contacts the bearing surfaces 98, 100, 102, 104 of the blade 38. A gap is defined between the sleeve 110 and the non-bearing surface of the lobes 90, 92 and another gap 113 is defined between the non-bearing surface of the lobes 94, 96.

In the second embodiment shown in Figure 3, the sleeve 110 is assembled to the blade 38 first. Then the blade 38 and sleeve 110 are inserted into the slot 72. The internal surface 116 of the sleeve 110 is in contact with the surfaces of the blade 38 comprising the radially inner surface 64 of the platform 64, the surface of the neck 88 and the surface of the root portion 84. A line 113 has a constant radius from the rotational axis 20 and passes through the radially outward edge of the bearing surfaces 98, 100 of the root portion 68. The inner surface 116 of the sleeve 110 is in contact with substantially all the surface 69 of the root portion 68 radially inwardly of the line 112. When the rotor assembly 70 is assembled the sleeve 110 is in contact with slot's bearing surfaces 76, 78, 80, 82 only.

It should be appreciated that the sleeve 110 may be assembled to the blade 38 (first) whether the root fixture 86 comprises a fir-tree configuration (multiple lobe design) or a single lobe dovetail design as described with reference to the first or the second embodiments respectively. Similarly, the sleeve 110 may be assembled to the slot 72 whether the slot 72 comprises a fir-tree configuration or a single lobe dovetail design as described with reference to the first or the second embodiments respectively.

The sleeve 110 is formed of a superalloy comprising a first composition having a Cr content % by weight and the blade 38 is formed of a superalloy comprising a second composition having a Cr content % by weight. The second composition has a Cr % by weight content that is greater than the first composition of the sleeve 110. In one example, the Cr content of the second composition is greater than 3% by weight Cr of the first composition. In another example, the sleeve 110 is formed from the first composition having a Cr content at least 12% by weight. The maximum Cr content may be less than 25% by weight. In one example, the first composition has a Cr content approximately 22% by weight. The blade 38 is formed from the second composition having a Cr content between 3% and 18% by weight.

Although preferably the second composition may have a Cr content less than 12% by weight.

Depending on the manufacturing technique used to form the sleeve 110, its superalloy composition may be dependant on commercially available alloys.

However, for the sleeve 110 formed by a casting process, one example of a superalloy composition has nominal constituents, Cr 22%, Co 19%, W 2%, Ta 1.1%, Al 2.3%, Ti 3.5%, Nb 0.8%, Hf 0.8%, Ni balance by weight.

The composition of the sleeve 110 and in particular the high-Cr content, gives the sleeve 110 a corrosion resistance and thereby protects the radially inner surface 64 of the platform 64 and the surface of the root portion 84. The longevity of the blade 38 is thereby not compromised by exposure to chemicals, particularly, sulphates which can corrode the blade 38 reducing it life. This corrosion is particularly severe radially inwardly of the upper surface 62 of the platform 64 where such corrosive chemicals can collect.

According to the present invention the sleeve 110 has a higher Cr content than the blade 38. This higher Cr content means that the sleeve 110 has a greater corrosion resistance than the material of the blade. Further, the lower Cr content of the blade means it has a greater strength and resistance to creep in service. In another example of the present invention, the Cr content is at least 3% by weight greater in the composition of the sleeve 110 than in the composition of the blade 38.

The sleeve 110 may be formed by pressing, an additive manufacturing technique or casting. It is possible to fabricate the sleeve 110 in two halves, split approximately at the intersection with the radial line 112 as shown in Figures 2 and 3, assemble the two halves to the blade or slot and then using a known joining technique join the two halves together.

In the case of the second embodiment, the sleeve 110 may be provided with cooling air inlets 118 that align with cooling passage 119 in the radially inner surface 117 of the blade 38.

In the method of assembling the rotor assembly, the sleeve 110 may be peened against the bearing surfaces 76, 78, 80, 82 of the slot 72 prior to inserting the root fixture 86 into the sleeve 110. Similarly, the method may comprise peening the sleeve 110 against the bearing surfaces 98, 100, 102, 104 of the root fixture 86 prior to inserting the root fixture 86 and the sleeve 110 into the slot 72. Furthermore, the sleeve 100 may be peened against all of the slot 72 in the first embodiment or against the surfaces of the blade the sleeve is in contact with in the second embodiment.

Referring to Figure 4, the sleeve 110 is shown in a perspective view. The internal surface 116 of the sleeve 110 can be seen with the blade 38 removed. The internal surface 116 would otherwise be in contact with the underside surface 64 of the platform 60. The sleeve 110 comprises a slot bearing surface 124 and a blade bearing surface 126. The slot bearing surface 124 and the blade bearing surface 126 are on opposite sides of the thickness 120 of the sleeve 110. At a leading end 128 of the sleeve 110 an end plate 127 is provided. The leading end 128 is the axially forward end of the blade and slot. The sleeve 110 is monolithically formed with the end plate 127, particularly where the sleeve 110 is manufactured by an additive manufacturing process or casting, or the end plate is integrally formed by welding or other joining technique if the sleeve is formed by pressing. The end plate 127 may be peened against the root portion of the blade if applied to the blade first as in the second embodiment.

Referring to Figure 5, where the sleeve 110 is applied to an existing rotor assembly 70 the slot 72 may require modifying such that a depth 120 of material is removed from its bearing surfaces 76, 78, 80, 82. The depth of material removed may be equivalent to the thickness of the sleeve 110 such that the tip of the blade 38 is in the same position as in the existing rotor assembly 70 without the sleeve 110.

Thus, the tip of the blade 38 remains in the same nominal position prior to application of the sleeve 110 and the tip gap is maintained. In one embodiment the sleeve 110 has a thickness of 3mm, but its thickness may be between 1mm and 4mm.

The superalloys referred to herein are any commercially available compositions such as Hastelloy, Inconel, Waspaloy, Rene alloys, Inc°loy, MP98T, TMS alloys, and CMSX single crystal alloys. Other superalloy compositions are intended to be included herein. Preferably, the superalloys are nickel-based superalloys, but other alloys are possible. The constituents of these superalloys may be adapted to comprising the prescribed Cr content % by weight.

Claims

CLAIMS1. A rotor assembly (70) for a gas turbine engine (10), the rotor assembly (70) comprises a rotational axis (20), a disc (36) and at least one blade (38), the disc (36) comprises at least one slot (72), the slot (72) has at least one pair of bearing surfaces (76, 78, 80, 82), the blade (38) comprises a root portion (84) having a root fixture (86) and radially outwardly a platform (60), the platform (60) has a radially inner surface (64), the root fixture (86) comprises at least one pair of bearing surfaces (98, 100, 102, 104), characterised in that a sleeve (110) is located at least partly between the slot (72) and the root fixture (86), the sleeve (110) is in contact with the slot's bearing surfaces (76, 78, 80, 82) and the root fixture's bearing surfaces (98, 100, 102, 104), the sleeve (110) extends and is in contact with at least a part of the radially inner surface (64).
2. The rotor assembly (70) as claimed in claim 1 wherein the root portion (84) comprises a neck (88), the neck (88) is between the root fixture (86) and the platform (60), the sleeve (110) is in contact with the neck (88).
3. The rotor assembly (70) as claimed in any one of claims 1-2 wherein the sleeve (110) is formed of a superalloy comprising a first composition having a Cr content % by weight, the blade (38) is formed of a superalloy comprising a second composition having a Cr content % by weight, wherein the second composition having a Cr content % by weight greater than the first composition having a Cr content % by weight.
The rotor assembly (70) as claimed in claim 3 wherein the second composition having a Cr content % by weight greater than 3% Cr the first composition having a Cr content % by weight.
5. The rotor assembly (70) as claimed in any one of claims 3-4 wherein the sleeve (110) is formed from the first composition having a Cr content at least 12% by weight, preferably having a Cr content less than 25% by weight, preferably Cr by weight, preferably having a Cr content approximately 22% by weight.
6. The rotor assembly (70) as claimed in any one of claims 3-5 wherein the blade (38) is formed from the second composition having a Cr content between 3% and 18% by weight, preferably the second composition having a Cr content less than 12%.
7. The rotor assembly (70) as claimed in any one of claims 1-6 wherein a line (112) has a constant radius from the rotational axis (20) and passes through the radially outward edge of the bearing surfaces (76, 78) of the slot (72), the sleeve (110) is in contact with substantially all the surface (73) of the slot (72) radially inwardly of the line (112).
8. The rotor assembly (70) as claimed in any one of claims 1-6 wherein a line (112) has a constant radius from the rotational axis (20) and passes through the radially outward edge of the bearing surfaces (98, 100) of the root portion (68), the sleeve (110) is in contact with substantially all the surface (69) of the root portion (68) radially inwardly of the line (112).
9. The rotor assembly (70) as claimed in any one of claims 1-8 wherein the root portion (68) comprises a radially inner surface (117) and a cooling passage (119) through the radially inner surface (117), the sleeve (110) comprises a cooling air inlet (118) that is aligned with cooling passage (119) to allow cooling air to pass into the root portion (68).
10. The rotor assembly (70) as claimed in any one of claims 1-9 wherein the sleeve (110) comprises an end plate (127), preferably the end plate (127) is at a leading end (128) of the sleeve (110).
11. A method of assembling a rotor assembly as claimed in any one of claims 1- 10, the method comprises inserting the sleeve (110) into the slot (72) and inserting the root fixture (86) into the sleeve (110).
12. A method of assembling a rotor assembly as claimed in claim 11, the method comprises peening the sleeve (110) against the bearing surfaces (76, 78, 80, 82) of the slot (72) prior to inserting the root fixture (86) into the sleeve (110).
13. A method of assembling a rotor assembly as claimed in any one of claims 1- 10, the method comprises inserting the root fixture (86) into the sleeve (110), inserting the root fixture (86) and the sleeve (110) into the slot (72).
14. A method of assembling a rotor assembly as claimed in claim 13, the method comprises peening the sleeve (110) against the bearing surfaces (98, 100, 102, 104) of the root fixture (86) prior to inserting the root fixture (86) and the sleeve (110) into the slot (72).
15. A method of manufacturing a sleeve (110) as claimed in any one of claims 1- 14, wherein the sleeve (110) is formed monolithically by a casting process or an additive manufacturing process.