ROTOR BLADE ARRANGEMENT FOR A TURBOMACHINE

Abstract

The present invention relates to a rotor blade arrangement for a turbomachine, with a rotor blade which has a sealing tip radially on the outside, and with a seal arrangement, wherein the seal arrangement forms a radially inwardly open cavity, in which the sealing tip is arranged, to which end the seal arrangement has a first sealing element, namely a first seal carrier with a first run-in coating, and a second sealing element, wherein the first run-in coating delimits the cavity radially on the outside, and the second sealing element delimits the cavity in an axial direction, and wherein the first and the second sealing element are assembled.

Claims

1.-15. (canceled)

16. A rotor blade arrangement for a turbomachine, wherein the arrangement comprises a rotor blade comprising a sealing tip radially on the outside and a sealing arrangement, the sealing arrangement forming a radially inwardly open cavity in which the sealing tip is arranged, to which end the sealing arrangement comprises a first sealing element, namely a first seal carrier with a first abradable liner, and a second sealing element, the first abradable liner delimiting the cavity radially on the outside and the second sealing element delimiting the cavity in an axial direction, and wherein the first sealing element and the second sealing element are assembled.

17. The rotor blade arrangement of claim 16, wherein a volume of the cavity≤2.Math.(H+S).Math.π.Math.R.Math.B, wherein R is a radius up to an outer wall surface of an outer shroud, H is a height of the sealing tip, B is a width of the outer shroud, and S is a gap width between the sealing tip and the first abradable liner.

18. The rotor blade arrangement of claim 17, wherein the volume of the cavity≤1.6.Math.(H+S).Math.π.Math.R.Math.B.

19. The rotor blade arrangement of claim 17, wherein the volume of the cavity≤1.2.Math.(H+S).Math.π.Math.R.Math.B.

20. The rotor blade arrangement of claim 16, wherein the second sealing element is a second seal carrier with a second abradable liner, the second abradable liner delimiting the cavity in an axial direction, and/or wherein the second sealing element is integrated into a guide vane adjacent downstream to the rotor blade or is connected directly to the latter, and/or the first sealing element is integrated into a guide vane adjacent upstream to the rotor blade or is connected directly to the latter.

21. The rotor blade arrangement of claim 20, wherein the second abradable liner has an axial overlap at at least one operating point with an outer shroud of the rotor blade on which the sealing tip projects radially outward.

22. The rotor blade arrangement of claim 21, wherein in a region of the axial overlap, a radial spacing between the second abradable liner and the outer shroud makes up no more than 3 times a radial thickness of the outer shroud.

23. The rotor blade arrangement of claim 21, wherein the second abradable liner has in each case an axial overlap with the outer shroud of the rotor blade at all operating points, taking into account a maximum possible axial offset of the outer shroud.

24. The rotor blade arrangement of claim 16, wherein the first seal carrier and/or a second seal carrier of the second sealing element delimits a hollow space with a rear side facing away from the cavity.

25. The rotor blade arrangement of claim 24, wherein a volume of the hollow space makes up at least 0.5 times a volume of the cavity.

26. The rotor blade arrangement of claim 16, wherein the first seal carrier forms the cavity with a first axial portion and radially delimits a gas duct of the turbomachine with a second axial portion.

27. The rotor blade arrangement of claim 26, wherein also a second seal carrier of the second sealing element forms the cavity with a first axial portion and radially delimits a gas duct of the turbomachine with a second axial portion.

28. The rotor blade arrangement of claim 27, wherein, viewed in each case in axial section, there is between axial portions of the first seal carrier a first bend point, and there is between axial portions of the second seal carrier a second bend point, wherein a reference point of an outer shroud, which is situated axially centrally in an inner wall surface, facing a gas duct, of the outer shroud, has at the most a radial spacing, from a connecting line between the first and the second bend point, which in terms of its amount makes up no more than 3 times a radial thickness of the outer shroud.

29. The rotor blade arrangement of claim 16, wherein the sealing arrangement comprises a third abradable liner which delimits the cavity in an opposite axial direction.

30. The rotor blade arrangement of claim 29, wherein the third abradable liner is arranged on the first seal carrier together with the first abradable liner.

31. The rotor blade arrangement of claim 29, wherein the third abradable liner has an axial overlap at at least one operating point with an outer shroud of the rotor blade on which the sealing tip projects radially outward.

32. A turbine module for a turbomachine, wherein the turbine module comprises the rotor blade arrangement of claim 16.

33. The turbine module of claim 32, wherein the rotor blade arrangement forms the last stage of the turbine module.

34. The turbine module of claim 32, wherein the second sealing element is arranged downstream from the first sealing element, relative to a flow around the rotor blade in a gas duct of the turbomachine.

35. A method for producing the rotor blade arrangement of claim 16, wherein the method comprises assembling the first and the second sealing element.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0031] The invention is explained in detail below with the aid of an exemplary embodiment, wherein the individual features in the subordinate claims can be essential to the invention also in other combinations and moreover no detailed distinction is made between the different categories of claim.

[0032] In detail, in the drawings:

[0033] FIG. 1 shows schematically a turbofan engine in an axial section;

[0034] FIG. 2 shows a rotor blade arrangement according to the invention in a schematic side view in partial section;

[0035] FIG. 3 shows a further rotor blade arrangement according to the invention in a schematic side view in partial section;

[0036] FIG. 4 shows the rotor blade arrangement according to FIG. 2 as an illustration of a radial situation;

[0037] FIG. 5 shows the rotor blade arrangement according to FIG. 2 as an illustration of an axial offset at different operating points.

PREFERRED EMBODIMENT OF THE INVENTION

[0038] FIG. 1 shows a turbomachine 1 in a schematic view, specifically a turbofan engine. The turbomachine 1 is divided functionally into a compressor 1a, a combustion chamber 1b, and a turbine 1c, the latter having a high-pressure turbine module 1ca and a low-pressure turbine module 1cb. Both the compressor 1a and the turbine 1c here each consist of multiple stages, and each stage is composed of a guide vane ring and a rotor blade ring. In each stage, the rotor blade ring is arranged downstream from the guide vane ring, relative to the flow around the gas duct 2. During operation, the rotor blades revolve about the longitudinal axis 3. The fan 4 is coupled via a gearbox 5 and the rotor blade rings of the low-pressure turbine module 1cb revolve faster than the fan 4 during operation.

[0039] FIG. 2 shows a rotor blade arrangement 20 with a rotor blade 21 which has a sealing tip 22 on the outside radially. The sealing tip 22 is arranged on an outer shroud 23, and the rotor blade leaf 24 is situated radially inside the latter. The rotor blade arrangement 20 moreover has a sealing arrangement 25 which forms a cavity 26 in which the sealing tip 22 is arranged. The sealing tip 22 is surrounded in the cavity 26 not only radially but also axially, which ensures a good sealing effect (cf the detailed introduction to the description).

[0040] The sealing arrangement 25 here has a multi-part structure, namely is composed of a first sealing element 31 and a second sealing element 32. This ensures axial ease of mounting, and when a turbine module 1ca,cb is mounted, the first sealing element 31 can first be mounted on the upstream guide vane 34 and the rotor blade 21 can then be positioned. The second sealing element 32 is then attached and the cavity 26 hence closed axially.

[0041] The first sealing element 31 has a first seal carrier 31.1 and a first abradable liner 31.2. In the present case, the seal carrier 31.1 is a curved plate and the abradable liner 31.2 is formed by a radially inwardly open cellular structure. The first abradable liner 31.2 delimits the cavity 26 radially on the outside.

[0042] The second sealing element 32 has a second seal carrier 32.1 and a second abradable liner 32.2, in the present case an axially open cellular structure. The second abradable liner 32.2 delimits the cavity 26 in an axial direction 35. There is moreover a third abradable liner 31.3 which is here arranged as part of the first sealing element 31 on the first seal carrier 31.1 and is designed as a single piece with the first abradable liner 31.2. The third abradable liner 31.3 delimits the cavity 26 in an axial direction 36 counter to the axial direction 35.

[0043] FIG. 3 shows a further rotor blade arrangement 20 according to the invention, wherein structurally identical parts of parts with the same function are provided with the same reference symbols and in this respect reference is made to the description of FIG. 2. In contrast to FIG. 2, the second sealing element 32 in this case does not have an abradable liner and instead the seal carrier 32.1 itself delimits the cavity 26 in the axial direction 35. This can represent a simplified variant in which, however, the sealing effect is also somewhat reduced.

[0044] FIG. 4 relates in turn to the rotor blade arrangement 20 according to FIG. 3. It can be seen therefrom initially that a first axial portion 41.1 of the first seal carrier 31.1 forms the cavity 26 and a second axial portion 41.2 of the first seal carrier 31.1 radially delimits the gas duct 2 of the turbomachine 1. Equally, the second seal carrier 32.1 forms the cavity 26 in a first axial portion 42.1 and it delimits the gas duct 2 radially in a second axial portion 42.2. The first seal carrier 31.1 has a first bend point 51 between the first and the second axial portion 41.1,41.2 and the second seal carrier 32.1 has a first bend point 52 between the first and the second axial portion 42.1,42.2. A reference point 23.1.1 of the outer shroud 23 is situated, relative to a connecting line 53, between the bend points 51,52, spaced apart by no more than 3 times a radial thickness 54 of the outer shroud 23. The reference point 23.1.1 is situated axially centrally in an inner wall surface 23.1 of the outer shroud 23.

[0045] FIG. 5 illustrates that the outer shroud 23 and hence the sealing tip 22 can assume different axial positions at different operating points, and there is therefore an axial offset 60. Two axial positions 61 are illustrated here in dotted lines which mark the maximum offset forward (for 61.1) and rearward (for 61.2). The operating points are the result of different operating states of the turbomachine, in the case of the aircraft engine for example at take-off and landing or when cruising (ADP). At least the second abradable liner 32.2 or the third abradable liner 31.3 preferably in each case have an axial overlap with the outer shroud 23, and particularly preferably both have an axial overlap.

LIST OF REFERENCE SIGNS

[0046] Turbomachine 1 [0047] Compressor 1a [0048] Combustion chamber 1b [0049] Turbine 1c [0050] High-pressure turbine module 1ca [0051] Low-pressure turbine module 1cb [0052] Gas duct 2 [0053] Longitudinal axis 3 [0054] Fan 4 [0055] Gearbox 5 [0056] Rotor blade arrangement 20 [0057] Rotor blade 21 [0058] Sealing tip 22 [0059] Outer shroud 23 [0060] Inner wall surface 23.1 [0061] Reference point 23.1.1 [0062] Rotor blade leaf 24 [0063] Sealing arrangement 25 [0064] Cavity 26 [0065] First sealing element 31 [0066] First seal carrier 31.1 [0067] First abradable liner 31.2 [0068] Third abradable liner 31.3 [0069] Second sealing element 32 [0070] Second seal carrier 32.1 [0071] Second abradable liner 32.2 [0072] Upstream guide vane 34 [0073] Axial direction 35 [0074] Opposite axial direction 36 [0075] Axial portions (first seal carrier) 41 [0076] First axial portion 41.1 [0077] Second axial portion 41.2 [0078] Axial portions (second seal carrier) 42 [0079] First axial portion 42.1 [0080] Second axial portion 42.2 [0081] First bend point 51 [0082] Second bend point 52 [0083] Connecting line 53 [0084] Radial thickness 54 [0085] Axial offset 60 [0086] Axial positions 61 [0087] Maximum forward offset 61.1 [0088] Maximum rearward offset 61.2 [0089] Width of outer shroud B [0090] Height of sealing tip H [0091] Radius (longitudinal axis to outer wall surface of outer shroud) R [0092] Gap width between sealing tip and first abradable liner S