# US 2013/0156563 A1 · Turbomachine and turbomachine stage

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- Number: US 2013/0156563 A1
- Document ID: US20130156563A1
- Status: Published application
- Kind: A1
- Application number: 13/687,798
- Filed: 2012-11-28
- Published: 2013-06-20
- Earliest claimed: 2011-12-20
- Assignees: MTU Aero Engines AG
- Inventors: Inga MAHLE; Jochen GIER; Kai KOERBER; Michael ENGELKE
- Web page: https://patentgrep.com/patent/US20130156563A1

## CPC

- F01D 5/141: {Shape, i.e. outer, aerodynamic form (F01D5/148 - F01D5/20 take precedence; blade construction F01D5/147)}
- F01D 5/143: {Contour of the outer or inner working fluid flow path wall, i.e. shroud or hub contour}
- F01D 5/225: {by shrouding}
- F01D 9/00: Stators
- F01D 9/04: forming ring or sector
- F01D 11/08: for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20)
- F04D 29/321: {for axial flow compressors}
- F04D 29/547: {having a special shape in order to influence fluid flow}
- F01D 11/001: {for sealing space between stator blade and rotor}
- F05D 2240/80: Platforms for stationary or moving blades
- F05D 2250/184: sinusoidal
- F05D 2250/611: undulated
- F05D 2250/70: Shape
- Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft
- F01D: NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES (machines or engines for liquids F03; non-positive displacement pumps F04D)
- F04D: NON-POSITIVE-DISPLACEMENT PUMPS (engine fuel-injection pumps F02M; ion pumps H01J41/12; electrodynamic pumps H02K44/02)
- F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION

## Abstract

A turbomachine stage includes guide vanes and an airfoil platform forming a guide vane cascade, and rotor blades and an airfoil platform forming a rotor blade cascade. Airfoil platforms have cascade regions extending between circumferentially adjacent airfoils, and gap regions which radially and/or axially bound an axial gap extending axially between the guide vane cascade and the rotor blade cascade. A contour of at least one of these gap regions varies in the radial and/or axial direction around the circumference. A maximum extent of this contour in the radial direction toward the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the radial direction being no more than 50% of the cascade pitch and/or a maximum extent in the axial direction away from the spoke-like pattern is circumferentially spaced from an airfoil edge of this cascade by no more than 50% of the cascade pitch, a maximum variation in the axial direction being no more than 50% of the cascade pitch.

## Claims

### Claim 1

A turbomachine stage comprising:
guide vanes and at least one of a radially inner and radially outer guide vane airfoil platform defining a guide vane cascade, the guide vane airfoil platform having a guide vane platform cascade region and a guide vane gap region; and
rotor blades and at least one of a radially inner and radially outer rotor blade airfoil platform defining a rotor blade cascade, the rotor vane airfoil platform having a rotor vane platform cascade region and a rotor vane gap region, the rotor blade cascade being adjacent to the guide vane cascade,
the guide vane gap region and the rotor vane gap region at least one of radially and axially bounding an axial gap extending axially between the guide vane cascade and the rotor blade cascade,
a contour of at least one of the guide vane and rotor vane gap regions varying in at least one of the radial and axial direction around a circumference,
at least one of (a) and (b) being true:
(a) a maximum extent of the contour in the radial direction toward a spoke-like pattern of the respective guide vane or rotor vane cascade being circumferentially spaced from a respective airfoil edge of the respective guide vane or rotor vane cascade by no more than 50% of the pitch of the respective guide vane or rotor vane cascade, a maximum variation of the contour in the radial direction being no more than 50%, of the pitch of the respective guide vane or rotor vane cascade; and
(b) a maximum extent of the contour in the axial direction away from the spoke-like pattern is circumferentially spaced from the respective airfoil edge of the respective guide or rotor vane cascade by no more than 50%, a maximum variation of the contour in the axial direction being no more than 50% of the pitch of the respective guide vane or rotor vane cascade.

### Claim 2 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies in the radial direction, and by a maximum extent of no more than 25%.

### Claim 3 (depends on 2)

The turbomachine stage as recited in claim 2 wherein the maximum variation in the radial direction is no more than 40%.

### Claim 4 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies in the radial direction and with the maximum variation in the radial direction being no more than 40%.

### Claim 5 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies in the axial direction, and by a maximum extent of no more than 25%.

### Claim 6 (depends on 5)

The turbomachine stage as recited in claim 5 wherein the maximum variation in the axial direction is no more than 40%.

### Claim 7 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies in the axial direction and with the maximum variation in the axial direction being no more than 40%.

### Claim 8 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the respective maximum extent is disposed in the pressure-side region of an airfoil leading edge or in the suction-side region of an airfoil trailing edge.

### Claim 9 (depends on 1)

The turbomachine stage as recited in claim 1 wherein an at least one of axially and radially opposite contour of the other of the guide vane and rotor blade cascades varies around the circumference.

### Claim 10 (depends on 9)

The turbomachine stage as recited in claim 9 wherein the opposite contour varies identically around the circumference as the contour.

### Claim 11 (depends on 9)

The turbomachine stage as recited in claim 9 wherein the opposite contour is radially opposite the contour.

### Claim 12 (depends on 11)

The turbomachine stage as recited in claim 11 wherein the opposite contour varies identically around the circumference as the contour.

### Claim 13 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies radially and axially.

### Claim 14 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies radially but is constant in the axial direction.

### Claim 15 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the respective guide vane platform or rotor vane platform gap region whose contour varies merges smoothly into the respective guide vane or rotor vane cascade region.

### Claim 16 (depends on 1)

The turbomachine stage as recited in claim 1 wherein the contour varies periodically.

### Claim 17 (depends on 1)

A turbomachine comprising at least one turbomachine stage as recited in claim 1.

### Claim 18 (depends on 1)

A gas turbine comprising at least one turbomachine stage as recited in claim 1.

### Claim 19 (depends on 1)

An aircraft engine gas turbine comprising at least one turbomachine stage as recited in claim 1.

### Claim 20 (depends on 1)

A compressor stage comprising the turbomachine stage as recited in claim 1.

### Claim 21 (depends on 1)

A turbine stage comprising the turbomachine stage as recited in claim 1.