Blade for a gas turbomachine

Patent metadata

Field	Value
Publication number	US-9896940-B2
Application number	US-201414325560-A
Country	US
Kind code	B2
Filing date	Jul 8, 2014
Priority date	Jul 9, 2013
Publication date	Feb 20, 2018
Grant date	Feb 20, 2018

Abstract

A blade, in particular a rotor blade or a stator vane, for a gas turbomachine, in particular a turbojet engine, the blade having an airfoil ( 1 ) for deflecting a flow of working fluid and a first platform ( 3 ) connected thereto, in particular integrally connected thereto, to radially bound a flow duct for the working fluid, the airfoil having a suction side and a pressure side ( 1.1 ) which are connected at a leading edge ( 1.2 ) and at a trailing edge ( 1.3 ). The trailing edge has a first minimum wall thickness in a first region (A) of a radial longitudinal extent (R) of the airfoil proximal to the first platform ( 3 ), and a maximum wall thickness that is smaller than the first minimum wall thickness in a platform-distal region (C) of the radial longitudinal extent.

First claim

What is claimed is:

1. A blade for a gas turbomachine, the blade comprising:
an airfoil for deflecting a flow of working fluid;
a first platform connected thereto to radially bound a flow duct for the working fluid,
the airfoil having a suction side and a pressure side connected at a leading edge and at a trailing edge, the trailing edge having a first minimum wall thickness in a first region of a radial longitudinal extent of the airfoil proximal to the first platform, and a maximum wall thickness smaller than the first minimum wall thickness in a platform-distal region of the radial longitudinal extent; and
a second platform radially opposite the first platform being connected to the airfoil to radially bound the flow duct, the trailing edge having a second minimum wall thickness greater than the maximum wall thickness in a second region of the radial longitudinal extent proximal to the second platform,
the maximum wall thickness of the trailing edge in the platform-distal region being no greater than 0.35 mm,
wherein each of the first and second platform proximal regions extends over 10% of the radial longitudinal extent in a direction away from a platform;
wherein the platform-distal region extends over at least 25% of the radial longitudinal extent from a middle of the radial longitudinal extent of the airfoil toward a platform-proximal region,
wherein the wall thickness of the trailing edge is equal to a maximum wall thickness of the airfoil in a region which extends by 5% from an axially rear end of the airfoil in the upstream direction toward the leading edge,
whereby the maximum wall thickness of the trailing edge in the platform distal region is at least substantially constant,
wherein the wall thickness of the trailing edge decreases strictly monotonically respectively, from the minimum wall thickness of the respective platform-proximal region to the maximum wall thickness of the platform-distal region in a transition region of the radial longitudinal extent between the first and second platform-proximal region and the platform-distal region, and
wherein the transition regions extend respectively over 10% of the radial longitudinal direction.

2. The blade as recited in claim 1 wherein the first or the second minimum wall thickness is at least 0.35 mm.

3. The blade as recited in claim 1 wherein the first or the second minimum wall thickness is at least 0.40 mm.

4. The blade as recited in claim 1 wherein the first or the second minimum wall thickness is at least 0.45 mm.

5. The blade as recited in claim 1 wherein the platform-distal region extends over at least 25 and no more than 30% of the radial longitudinal extent from a middle of the radial longitudinal extent of the airfoil toward the first or the second platform.

6. The blade as recited in claim 1 wherein the trailing edge of the airfoil merges into a platform in a corner.

7. The blade as recited in claim 6 wherein the corner is concavely curved.

8. A rotor blade comprising the blade as recited in claim 1 .

9. A stator vane comprising the blade as recited in claim 1 .

10. A turbojet engine comprising the blade as recited in claim 1 .

11. The blade as recited in claim 1 wherein the airfoil and the first platform are integrally connected.

12. The blade as recited in claim 1 wherein the airfoil and the second platform are integrally connected.

13. A gas turbomachine comprising at least one compressor stage or at least one turbine stage having at least one blade as recited in claim 1 , the at least one blade being detachably or permanently connected to a rotor of the gas turbomachine.

14. The gas turbomachine as recited in claim 13 wherein the blade is integrally connected to the rotor.

15. A turbojet engine comprising the gas turbomachine as recited in claim 13 .

Assignees

• MTU Aero Engines AG

Inventors

• Boeck Alexander
• Huebner Norbert
• Schmid Sven

Classifications

• F01D5/147 Primary

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title

• F01D5/225

  by shrouding · CPC title

• F01D5/141

  Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147) · CPC title

• F05D2240/304

  related to the trailing edge of a rotor blade · CPC title