US20130266427A1

US20130266427A1 - Sealing system for a turbomachine

Info

Publication number: US20130266427A1
Application number: US13/856,535
Authority: US
Inventors: Inga Mahle; Alexander Boeck
Original assignee: MTU Aero Engines GmbH
Current assignee: MTU Aero Engines AG
Priority date: 2012-04-04
Filing date: 2013-04-04
Publication date: 2013-10-10
Also published as: EP2647796A1

Abstract

A sealing system for a turbomachine, in particular, for a gas turbine, having an annular space between a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades. The sealing system comprises at least one first sealing fin disposed on an end of the rotor blade facing the wall, at least one second abradable lining disposed following the first sealing fin in the flow direction on the end of the rotor blade facing the wall, at least one first abradable lining disposed on the inside of the wall and opposite the first sealing fin, and at least one second sealing fin disposed in the flow direction following the first abradable lining on the inside of the wall and opposite the second abradable lining. The invention furthermore relates to a gas turbine comprising at least one sealing system.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS

This application claims benefit of and priority to European Patent Application No. 12163065.1, filed Apr. 4, 2012, entitled “Dichtungssystem für eine Strömungsmaschine,” the specification of which is incorporated herein in its entirety.

TECHNICAL FIELD

The invention relates to a sealing system for a turbomachine, in particular, for a gas turbine, wherein the sealing system is disposed in an annular space between a flow-restricting wall of the turbomachine and at least one series of rotor blades comprising a plurality of rotor blades, and at least two seals. The invention furthermore relates to a gas turbine, in particular, an aircraft engine that includes at least one sealing system.

BACKGROUND

Sealing systems of this type are employed, in particular, in connection with so-called clearance-maintaining systems in the components of compressors and turbines. These sealing systems function to keep to a minimum a sealing gap of the rotating blading relative to a housing, as well as the clearances of a stationary blading relative to a rotating rotor hub, thereby ensuring stable operating characteristics together with high efficiency. The rotating components of the turbine typically include sealing fins that in the known way touch or run in against honeycomb seals. The seals are provided here in the form of rub linings and abradable linings Sealing systems of this type are disclosed, for example, in U.S. Pat. No. 4,856,963 B1 and DE 198 07 247 A1. By minimizing the radial gaps above the sealing fins, an attempt is made to minimize the leakage flows through cavities in these regions, in particular, in the regions above the shrouds of the rotor blades, and the efficiency losses generated thereby. Nevertheless, mixing losses are generated when the leakage flows enter the so-called main flow of the turbomachine due to the different orientations and velocities of the main flow and the leakage flow. In addition, the flow around a following blade is not optimal and this results in further losses.

SUMMARY AND DESCRIPTION

The object of this invention is therefore to create a sealing system of the type referenced above that reliably provides an increase in the efficiency of a turbomachine. A further object of the invention is to provide a corresponding gas turbine that exhibits improved efficiency.
The objects of the invention are achieved according to the invention by a sealing system having the features described and claimed herein and a gas turbine having the features described and claimed herein. A dvantageous embodiments of the invention are provided in the respective subordinate claims, wherein advantageous embodiments of the sealing system must be considered as advantageous embodiments of the gas turbine according to the invention. Similarly, advantageous embodiments of the gas turbine according to the invention must be considered as advantageous embodiments of the sealing system.
A sealing system according to the invention for a turbomachine, in particular, for a gas turbine, is disposed in an annular space between a flow-restricting wall of the turbomachine and at least one series of rotor blades comprising a plurality of rotor blades. The sealing system here comprises at least two seals, wherein the first seal comprises at least one first sealing fin that is disposed on an end of the rotor blade facing the wall as well as at least one first abradable lining that is disposed on an inside of the wall and that is opposite the first sealing fin, and the second seal is disposed in the direction of flow following the first seal, and comprises a second abradable lining that is provided on the end of the rotor blade facing the wall as well as a second sealing fin that is disposed on the inside of the wall and is opposite the second abradable lining. The first sealing fin and the second abradable lining here can be disposed on a shroud of the rotor blade. In terms of the flow-restricting wall, this can be a casing of the turbomachine. The sealing system according to the invention is able to ensure that any leakage flow is conducted closely along the end of the rotor blade facing the wall or along the shroud of the rotor blade until it reenters the so-called main flow of the turbomachine, with the result that this is delayed less severely until it reenters the main flow as compared with conventional sealing systems, in particular, sealing systems that have at least two sealing fins on the shroud of the rotor blade. In addition, the leakage flow can be introduced with less disruption, that is, with less generation of turbulence, into the main flow. This results in an overall strong reduction in the velocity gradients between the leakage flow and the main flow, thereby reducing the mixing losses and significantly improving the efficiency of the turbomachine. In addition, this results in a more advantageous flow around the following blade. The seal of the sealing system according to the invention, which seal is disposed last in the flow direction, is always of the same design as the referenced second seal in order to achieve the described advantages. Otherwise the sealing system according to the invention can be implemented using already known techniques. The sealing system can thus be of relatively lightweight construction so as to prevent any failure of the bond between the second abradable lining and the rotor blade from causing the entire sealing system to fail since those components of the sealing system that are located further forward in the flow direction, that is, the first seal, continue to be present. Also minimized is any secondary damage, such as, for example, that caused by the components of the second abradable lining impacting the stages of the turbomachine located further back in the flow direction, due to the relatively small size of the second abradable lining.
In further advantageous embodiments of the sealing system according to the invention, the second abradable lining has at least one chamfer in a rearward region in the flow direction. However, it is also possible for the shroud to have at least one chamfer in a rearward region in the flow direction. The chamfer here can be at an angle a ranging between 10° and 65°, preferably, approximately 20° and 45°, relative to a longitudinal axis of the turbomachine. This design shape of the second abradable lining and/or of the shroud enables the generation of turbulence to be further reduced in the leakage flow in these regions, and this results in a significantly more trouble-free entrance of the leakage flow into the main flow of the turbomachine. This in turn provides a significant reduction in the velocity gradients between the leakage flow and the main flow, thereby reducing mixing losses and improving the efficiency of the turbomachine.
In another advantageous embodiment of the sealing system according to the invention, the first and/or second sealing fin is tilted opposite the flow direction. This also enables turbulences to be significantly reduced within the leakage flow.
In further advantageous embodiments of the sealing system according to the invention, the first sealing fin is integrated into the shroud. It is also possible to integrate the second sealing fin in the flow-restricting wall. This yields production advantages that result in overall reduced manufacturing costs.
In another advantageous embodiment of the sealing system according to the invention, the sealing system is provided in the form of a stepped labyrinth. This allows the sealing system to be advantageously adapted to the relevant specific requirements of the turbomachine, in particular, to the requirements presented by a low-pressure turbine.
The invention furthermore relates to a gas turbine, in particular, an aircraft engine, comprising at least one sealing system, wherein the sealing system is disposed in an annular space between a flow-restricting wall of the gas turbine and at least one series of rotor blades comprising a plurality of rotor blades, and at least two seals, wherein the first seal comprises at least one sealing fin that is disposed on an end of the rotor blade facing the wall as well as at least one first abradable lining that is disposed on an inside of the wall and is opposite the first sealing fin, and the second seal is disposed in the flow direction following the first seal and comprises a second abradable lining that is provided on the end of the rotor blade facing the wall as well as a second sealing fin that is disposed on the inside of the wall and is opposite the second abradable lining. The first sealing fin and the second abradable lining here can be disposed on a shroud of the rotor blade. The flow-restricting wall can be a casing of the turbomachine. The gas turbine according to the invention exhibits improved efficiency due to the fact that a leakage flow can be conducted by the sealing system closely along the end of the rotor blades facing the flow-restricting wall up to re-entry into a main flow of the gas turbine such that this maintains a high velocity up to the point of re-entry, with the result that the velocity difference vis-à-vis the relatively fast main flow is reduced. In addition, the referenced sealing system ensures low generation of turbulence in the leakage flow, thereby allowing this flow to be introduced into the main flow in relatively trouble-free fashion. In overall terms, this results in a reduction in the referenced mixing losses and in a more advantageous flow around the following blade. In order to achieve these advantages, the seal that is disposed last in the flow direction is always of the same design as the referenced second seal. The sealing system for the gas turbine according to the invention can otherwise be implemented using already known techniques. The sealing system can thus be of relatively lightweight construction so that any failure of the bond between the second abradable lining and the rotor blade does not result in failure of the entire sealing system due to the fact that the components of the sealing system located further forward in the flow direction continue to be present. In addition, any consequential damage, such as, for example, that caused by components of the second abradable lining impacting stages of the turbomachine located further back in the flow direction are minimized due to the relatively small size of the second abradable lining. Also minimized is the risk of a rotor imbalance in the turbomachine caused by asymmetric rubbing in the gas turbine according to the invention, or by the sealing system according to the invention.
Additional embodiments of the sealing system of the gas turbine have been described above.
Additional features of the invention are revealed in the claims, the exemplary embodiment, and based on the drawing. The features and combinations of features referenced above in the description, as well as the features and combinations of features referenced below in the exemplary embodiment, can be applied not only in the specific combination indicated but also in other combinations without falling outside the scope of the invention.

BRIEF DESCRIPTION OF THE DRAWING

The FIGURE here provides a sectional, schematic, and partial cutaway view of a sealing system according to the invention.

DETAILED DESCRIPTION

The sealing system 10 depicted in the figure is a component of a low-pressure gas turbine. It is evident that sealing system 10 is disposed in an annular space 30 between a flow-restricting wall 22 of the low-pressure gas turbine and a series of rotor blades comprising a plurality of rotor blades 12. Sealing system 10 here comprises two seals 34, 36, second seal 36 being disposed following first seal 34 in the flow direction 28. First seal 34 comprises a first sealing fin 18 that is provided on a shroud 16 of rotor blade 12, and a first abradable lining 24 that is disposed on an inside of wall 22 and is opposite first sealing fin 18. Second seal 36 comprises a second abradable lining 20 that is disposed following first sealing fin 18 in flow direction 28, and a second sealing fin 26 that is disposed on the inside of wall 22, which sealing fin is disposed following first abradable lining 24 in flow direction 28 and is opposite second abradable lining 20. First and second rub and abradable linings 24, 20 can be provided in the conventional form. They have a honeycomb structure in the embodiment shown. Another possible approach is to produce first and/or second abradable linings 24, 20 by spraying them onto shroud 16 or the inside of wall 22. It is possible, in particular, to employ powder injection molding for this purpose. It is also possible, however, to dispose first and/or second abradable linings 24, 20 as a complete system in the referenced regions or to bond them to these regions, for example, by brazing. Second abradable lining 20 in the illustrated embodiment is disposed axially following a reinforcement rib 38 that is created by a Z-interlocking of shrouds 16. This advantageously creates a flat attachment surface for second abradable lining 20 on the surface of shroud 16.
It is furthermore evident that second abradable lining 20 has at least one chamfer 32 in rearward region in flow direction 28. Chamfer 32 in the illustrated embodiment here is at an angle a less than 30° relative to a longitudinal axis of the turbomachine, for example, of the low-pressure turbine. It is possible in general for angle to lie within a range between 10° and 65°, preferably, a range between 20° and 45°.
It is furthermore evident that depicted sealing system 10 is provided in the form of a stepped labyrinth. First and second abradable linings 24, 20, as well as first and second sealing fins 18, 26 of sealing system 10 can be composed of such materials as are typically used. A wide variety of materials is known that can be used here. First sealing fin 18 in the illustrated embodiment is integrated into shroud 16. Second sealing fin 26 is attached to the inside of wall 22. In the illustrated embodiment, a guide vane 14 is disposed following rotor blade 12 in the flow direction.
The sealing system described in the embodiment is not restricted to the area of low-pressure turbines.

Claims

What is claimed is:

1. A sealing system for a turbomachine having a flow-restricting wall and at least one series of rotor blades including a plurality of rotor blades, the flow-restricting wall and the at least one series of rotor blades defining an annular space therebetween having a flow direction, the sealing system comprising:

a first seal and a second seal disposed in the annular space between the flow-restricting wall and the at least one series of rotor blades;

wherein the first seal includes

a first sealing fin that is disposed on an end of a rotor blade facing the flow-restricting wall, the rotor blade being from the at least one series of rotor blades, and

a first abradable lining that is disposed on an inside of the flow-restricting wall and is opposite the first sealing fin; and

wherein the second seal is disposed following the first seal in the flow direction and includes

a second abradable lining that is disposed on the end of the rotor blade facing the flow-restricting wall, and

a second sealing fin that is disposed on the inside of the flow-restricting wall and is opposite the second abradable lining.

2. A sealing system in accordance with claim 1, wherein the first sealing fin and the second abradable lining are disposed on a shroud of the rotor blade.

3. A sealing system in accordance with claim 2, wherein the shroud has at least one chamfer in a rearward region in the flow direction.

4. A sealing system in accordance with claim 3, wherein the chamfer is at an angle a in the range between 10° and 65° relative to a longitudinal axis of the turbomachine.

5. A sealing system in accordance with claim 4, wherein the chamfer is at an angle a in the range between 20° and 45° relative to a longitudinal axis of the turbomachine.

6. Sealing system according to claim 2, characterized in that the first sealing fin is integrated into the shroud.

7. A sealing system in accordance with claim 1, wherein the second abradable lining has at least one chamfer in a rearward region in the flow direction.

8. A sealing system in accordance with claim 7, wherein the chamfer is at an angle a in the range between 10° and 65° relative to a longitudinal axis of the turbomachine.

9. A sealing system in accordance with claim 8, wherein the chamfer is at an angle a in the range between 20° and 45° relative to a longitudinal axis of the turbomachine.

10. A sealing system in accordance with claim 1, wherein at least one of the first sealing fin or the second sealing fin is tilted opposite the flow direction.

11. A sealing system in accordance with claim 1, wherein the second sealing fin is integrated into the flow-restricting wall.

12. A sealing system in accordance with claim 1, wherein the sealing system is provided in the form of a stepped labyrinth.

13. A gas turbine aircraft engine comprising:

a casing wall;

at least one series of rotor blades including a plurality of rotor blades, the casing wall and the at least one series of rotor blades defining an annular space therebetween having a flow direction;

a first seal disposed in the annular space between the casing wall and the at least one series of rotor blades;

the first seal including a first sealing fin that is disposed on an end of a rotor blade facing the casing wall, the rotor blade being from the at least one series of rotor blades, and

a first abradable lining that is disposed on an inside of the casing wall and is opposite the first sealing fin; and

a second seal disposed in the annular space following the first seal in the flow direction and including

a second abradable lining that is disposed on the end of the rotor blade facing the casing wall, and

a second sealing fin that is disposed on the inside of the casing wall and is opposite the second abradable lining.

14. A gas turbine aircraft engine in accordance with claim 13, wherein the first sealing fin and the second abradable lining are disposed on a shroud of the rotor blade.

15. A gas turbine aircraft engine in accordance with claim 14, wherein the first sealing fin is integrated into the shroud.

16. A gas turbine aircraft engine in accordance with claim 13, wherein at least one of the first sealing fin or the second sealing fin is tilted opposite the flow direction.

17. A gas turbine aircraft engine in accordance with claim 13, wherein the second sealing fin is integrated into the casing wall.