METHOD FOR MANUFACTURING A TURBINE ASSEMBLY

Abstract

A method for manufacturing a turbine assembly having at least one aerofoil unit including at least a basically hollow aerofoil with at least one cooling passage for a cooling medium and at least one entry surface, wherein the at least one cooling passage enters at the at least one entry surface, and further the turbine assembly has at least one cover plate that at least partially covers the at least one entry surface. In order to provide a reliable attachment the method includes the step of attaching the at least one cover plate with one single, continuous, connecting structure to the at least one aerofoil unit.

Claims

1. A method for manufacturing a turbine assembly comprising at least one aerofoil unit comprising at least a basically hollow aerofoil with at least one cooling passage for a cooling medium and at least one entry surface, wherein the at least one cooling passage enters at the at least one entry surface, and further the turbine assembly comprises at least one cover plate that at least partially covers the at least one entry surface, wherein the method comprises: attaching the at least one cover plate with one single, continuous, connecting structure to the at least one aerofoil unit.

2. The method according to claim 1, wherein the at least one cover plate comprises at least one end, a centroid and an edge point, wherein a metric function of the centroid and the edge point has a maximum and wherein the maximum is located at the at least one end of the at least one cover plate and wherein the method comprises: attaching the at least one cover plate in such a way so that the at least one end is a free end in respect to the at least one aerofoil unit.

3. The method according to claim 1, wherein the at least one cover plate comprises a centroid and wherein the method comprises: attaching the at least one cover plate in such a way that the connecting structure extends through the centroid.

4. The method according to claim 1, wherein the method comprises: attaching the at least one cover plate by welding.

5. The method according to claim 1, wherein the at least one cover plate comprises at least two orifices that are in communication with the cooling passage wherein the method comprises: attaching the at least one cover plate to the at least one aerofoil unit basically in a middle between the at least two orifices of the at least one cover plate.

6. A turbine assembly comprising at least one aerofoil unit comprising at least one basically hollow aerofoil with at least one cooling passage for a cooling medium and at least one entry surface, wherein the at least one cooling passage enters at the at least one entry surface, and further comprising at least one cover plate that at least partially covers the at least one entry surface, wherein the turbine assembly is manufactured according to the method of claim 1.

7. The turbine assembly according to claim 6, wherein the at least one aerofoil unit comprises at least two apertures communicating with the at least one cooling passage and wherein the connecting structure extends through a mid-point being located basically in a middle between the at least two apertures.

8. The turbine assembly according to claim 6, wherein the at least one cover plate comprises at least one border, a centroid and an edge point, wherein a metric function of the centroid and the edge point has a maximum and wherein a maximum of the metric function is located on the at least one border of the at least one cover plate and wherein the at least one border is unattached to the at least one aerofoil unit.

9. The turbine assembly according to claim 6, wherein the at least one cover plate comprises two opposed borders, wherein the connecting structure extends between the two opposed borders.

10. The turbine assembly according to claim 6, wherein the at least one cover plate comprises two opposed borders, wherein the connecting structure extends all between the two opposed borders.

11. The turbine assembly according to claim 6, wherein the at least one cover plate comprises a first set of two opposed borders and a second set of two opposed borders, wherein the first set of two opposed borders are shorter than the second set of two opposed borders and/or wherein the connecting structure extends basically perpendicular to the opposed borders of the second set of two opposed borders.

12. The turbine assembly according to claim 6, wherein the at least one cover plate has a basically tetragonal shape and preferably, or a basically rectangular shape.

13. The turbine assembly according to claim 6, wherein the at least one aerofoil unit comprises at least two apertures communicating with the at least one cooling passage and wherein the at least one cover plate comprises at least one orifice communicating with at least one aperture of the at least two apertures of the aerofoil unit.

14. The turbine assembly according to claim 6, wherein the at least one aerofoil unit comprises at least two apertures communicating with the at least one cooling passage and wherein the at least one cover plate comprises at least one orifice that has a smaller diameter (d) than a diameter (D) of at least one aperture of the at least two apertures of the aerofoil unit.

15. A method of sealing using Use of a cover plate as a sealing plate, the method comprising: sealing with the cover plate at least one cooling passage of an aerofoil unit of a turbine assembly according to claim 6 to prevent a flow of cooling medium into and/or from the at least one cooling passage during operation of the turbine assembly.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0043] The present invention will be described with reference to drawings in which:

[0044] FIG. 1: shows a schematically and sectional view of a gas turbine engine comprising several inventive turbine assemblies,

[0045] FIG. 2: shows a perspective view of one turbine assembly of FIG. 1 with an aerofoil unit with a cut-away section showing cooling passages and a cover plate,

[0046] FIG. 3: shows a cross section through a root portion of the turbine assembly along line III-III in FIG. 2,

[0047] FIG. 4: shows a bottom view of the turbine assembly showing the connecting structure attaching the cover plate to the root portion,

[0048] FIG. 5: shows a bottom view of the turbine assembly with an entry surface for the cooling passage shown in FIG. 2 and

[0049] FIG. 6: shows the cover plate from FIG. 2 unattached to the aerofoil unit.

DETAILED DESCRIPTION OF THE ILLUSTRATED EMBODIMENTS

[0050] The terms upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine 44 unless otherwise stated. If used and not otherwise stated, the terms axial, radial and circumferential are made with reference to a rotational axis 54 of the engine 44.

[0051] FIG. 1 shows an example of a gas turbine engine 44 in a sectional view. The gas turbine engine 44 comprises, in flow series, an inlet 46, a compressor section 48, a combustion section 50 and a turbine section 52, which are generally arranged in flow series and generally in the direction of a longitudinal or rotational axis 54. The gas turbine engine 44 further comprises a shaft 56 which is rotatable about the rotational axis 54 and which extends longitudinally through the gas turbine engine 44. The shaft 56 drivingly connects the turbine section 52 to the compressor section 48.

[0052] In operation of the gas turbine engine 44, air 58, which is taken in through the air inlet 46 is compressed by the compressor section 48 and delivered to the combustion section or burner section 50. The burner section 50 comprises a burner plenum 60, one or more combustion chambers 62 defined by a double wall can 64 (not shown in detail) and at least one burner 66 fixed to each combustion chamber 62. The combustion chamber(s) 62 and the burner(s) 66 are located inside the burner plenum 60. The compressed air passing through the compressor section 48 enters a diffuser 68 and is discharged from the diffuser 68 into the burner plenum 60 from where a portion of the air enters the burner 66 and is mixed with a gaseous or liquid fuel. The air/fuel mixture is then burned and the combustion gas 70 or working gas from the combustion is channelled via a transition duct 72 to the turbine section 52.

[0053] The turbine section 52 comprises a number of blade carrying production discs 74 or turbine wheels attached to the shaft 56. In the present example, the turbine section 52 comprises four discs 74 each carry an annular array of turbine assemblies 10 which each comprises an aerofoil unit 12 (see FIG. 2) with an aerofoil 14 embodied as a turbine blade. However, the number of blade carrying production discs 74 could be different, i.e. only one production disc 74 or more than four production discs 74. In addition, turbine assemblies 10 embodied as turbine cascades 76 with aerofoil units 12 are disposed between the turbine blades 42. Each turbine cascade 76 carries an annular array of aerofoil units 12 which each comprises an aerofoil 14 in the form of guiding vanes, which are fixed to a stator 78 of the gas turbine engine 44. Between the exit of the combustion chamber 62 and the leading turbine blades inlet guiding vanes or nozzle guide vanes 80 are provided.

[0054] The combustion gas 70 from the combustion chamber 62 enters the turbine section 52 and drives the turbine blades which in turn rotate the shaft 56. The guiding vanes 80 serve to optimise the angle of the combustion or working gas 70 on to the turbine blades. The compressor section 48 comprises an axial series of guide vane stages 82 and rotor blade stages 84 with turbine assemblies 10 comprising aerofoil units 12 or turbine blades or vanes, respectively. In circumferential direction 86 around the turbine assemblies 10 the turbine engine 44 comprises a stationary casing 88.

[0055] FIG. 2 shows in a perspective view a turbine assembly 10 of the gas turbine engine 44 with an aerofoil unit 12 and a cover plate 22. The aerofoil unit 12 comprises a basically hollow aerofoil 14 embodied as a turbine blade, with two cooling regions, specifically, a channelled cooling region 90 and a fin-pin/pedestal cooling region 92. The former is located at a leading edge 94 and the latter at a trailing edge 96 of the aerofoil 14. The aerofoil 14 or its channelled cooling region 90, respectively, comprises two cooling passages 16 for a cooling medium 18. The cooling passages 16 extend in span wise direction 98 of the aerofoil 14 and are separated by rips 100. Moreover, the cooling passages 16 may be in flow communication with each other or with other cooling features of the aerofoil 14, like film cooling holes, impingement devices or the like (not specified or shown).

[0056] The aerofoil unit 12 further comprises a platform 102 and a root portion 104, wherein the platform 102 is arranged in span wise direction 98 between the aerofoil 14 and the root portion 104. Further, the aerofoil unit 12 may comprise an outer platform, embodied as a shroud, at its tip, which is not shown in FIG. 2. Moreover, the aerofoil is basically sealed at its tip. The aerofoil unit 12 or its root portion 104 comprises an entry surface 20, wherein the cooling medium 18 enter at the entry surface 20 through apertures 36 communicating with the cooling passages 16 (see FIG. 3 that shows a cross section along line III-III in FIG. 2). The cover plate 22 is attached to the aerofoil unit 12 or its root portion 104 at the entry surface 20 to partially cover the entry surface 20 or the apertures 36.

[0057] This could be seen in FIG. 4 that shows a bottom view of the turbine assembly 10 with the attached cover plate 22, wherein the covered apertures 36 of the root portion 104 are shown in dashed lines. The cover plate 22 has a basically tetragonal and rectangular shape and in this exemplary embodiment two orifices 32 communicating with the two apertures 36 of the aerofoil unit 12 or the cooling passages 16. Each orifice 32 of the cover plate 22 has a smaller diameter d than a diameter D of the two apertures 36 of the aerofoil unit 12.

[0058] The cover plate 22 is attaching to the aerofoil unit 12 or its root portion 104, respectively, with one single, continuous, connecting structure 24. This is done by welding the cover plate 22 to the aerofoil unit 12, thus forming one single, continuous, one-directional weld.

[0059] The connecting structure 24 or the weld, respectively, is positioned basically in a middle 34 between the two orifices 32 of cover plate 22 (see also FIG. 6 that shows the cover plate unattached to the aerofoil unit 12) and extends through a mid-point 38 being located basically in a middle between the two apertures 36 of the root portion 104 (see also FIG. 5 that shows a bottom view of the turbine assembly 10 with the entry surface 20). Thus, the orifices 32 of the cover plate 22 and the apertures 36 of the root portion 104 are arranged aligned to each other and with mirror symmetry to each other, wherein the connecting structure 24 is the symmetry axis.

[0060] Due to the rectangular shape of the cover plate 22 it comprises four borders 40, 40′, 42, 42′, wherein two sets of opposed borders 40, 40′, 42, 42′ or a first set of two opposed borders 40, 40′ and a second set of two opposed borders 42, 42′ are formed. The first set of two opposed borders 40, 40′ are shorter than the second set of two opposed borders 42, 42′. The connecting structure 24 extends all between two opposed borders 42, 42′ and specifically between the second set of longer borders 42, 42′ and basically perpendicular the second set of two opposed borders 42, 42′. In other words the connecting structure 24 may extend from one border 42 (40) to the other opposing border 42′ (42); alternatively, the connecting structure 24 may extend only part of the distance between the opposed borders 42, 42′ or 40, 42′ and not necessarily from one or to the other opposed border. The connecting structure 24 may be arranged to run along the longest dimension of the cover plate 22 or along the shortest dimension of the cover plate 22 as shown in FIG. 4 for example. The connecting structure 24 may be generally perpendicular to the borders 42, 42′ as shown in FIG. 4, but the connecting structure may be angled from perpendicular to one or both borders.

[0061] Moreover, the cover plate 22 has a centroid 28 and the connecting structure 24 extends through the centroid 28 (see also FIG. 5). Further, the cover plate 22 has several edge points 30 at one end 26, 26′ or as a part of one border 40, 40′ of the cover plate 22 (for better presentability only one edge point 30 is marked for each end 26, 26′/border 40, 40′). A metric function of the centroid 28 and each edge point 30 has a maximum and the maximum is or all maxima are located at the end 26, 26′ or the border 40, 40′.

[0062] By attaching the cover plate 22 via the connecting structure 24 that extends through the centroid 28 the ends 26, 26′ or the borders 42, 42′ with the edge points 30 are free or unattached to the aerofoil unit 12 or the root portion 104. Thus, in the attached state the cover plate 22 has free ends 26, 26′. During operation of the engine 44 and the turbine assembly 10 centrifugal forces are acting on the cover plate 22. Due to the free ends 26, 26′ the cover plate 22 is or the ends 26, 26′ are able to be pressed tight to the entry surface 20. Hence, the cover plate 22 seals the cooling passages 16 to prevent an unintended flow of cooling medium 18 into and/or from the cooling passage 16 during operation of the turbine assembly 10. Consequently, the cover plate 22 is used as a sealing plate.

[0063] It should be noted that the term “comprising” does not exclude other elements or steps and “a” or “an” does not exclude a plurality. Also elements described in association with different embodiments may be combined. It should also be noted that reference signs in the claims should not be construed as limiting the scope of the claims.

[0064] Although the invention is illustrated and described in detail by the preferred embodiments, the invention is not limited by the examples disclosed, and other variations can be derived therefrom by a person skilled in the art without departing from the scope of the invention.