Turbomachine and turbomachine stage

Patent metadata

Field	Value
Publication number	US-9863251-B2
Application number	US-201213687798-A
Country	US
Kind code	B2
Filing date	Nov 28, 2012
Priority date	Dec 20, 2011
Publication date	Jan 9, 2018
Grant date	Jan 9, 2018

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.

First claim

What is claimed is:

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 blade airfoil platform having a rotor blade platform cascade region and a rotor blade gap region, the rotor blade cascade being adjacent to the guide vane cascade,
the guide vane gap region and the rotor blade 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 blade 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 blade cascade being circumferentially spaced from a respective airfoil edge of the respective guide vane or rotor blade cascade by no more than 50% of the pitch of the respective guide vane or rotor blade 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 blade 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 blade 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 axial direction being no more than 50% of the pitch of the respective guide vane or rotor blade cascade;
and
whereby a radially opposite contour of the other of the guide vane and the rotor blade gap regions varies identically around the circumference as the contour but with a phase shift, wherein:
the contour varies according to R(φ)=R 0 +ΔR×sin(Ω R ×φ) and the radially opposite contour varies with the phase shift PS(x) according to R(φ, x)=R 0 +ΔR×sin(Ω R ×φ+PS(x)), wherein the phase shift PS(x) varies with the axial position x,
the contour varies according to X(φ)=X 0 +ΔX×sin(Ω x ×φ) and the radially opposite contour varies with the phase shift PS(r) according to X(φ, r)=X 0 +ΔX×sin(Ω x ×φ+PS(r)), wherein the phase shift PS(r) varies with the radial position r,
wherein φε[0°, 360°], R 0 , ΔR, X 0 , ΔX, Ω R , Ω x , Φ R , Φ x are constant, R defines a variation in a radial direction, and X defines a variation in an axial direction.

2. The turbomachine stage as recited in claim 1 wherein the contour varies in the radial direction, the maximum extent of the contour in the radial direction toward the spoke-like pattern of the respective guide vane or rotor blade cascade being circumferentially spaced from a respective airfoil edge of the respective guide vane or rotor blade cascade by no more than 25% of the pitch of the respective guide vane or rotor blade cascade.

3. The turbomachine stage as recited in claim 2 wherein the maximum variation in the radial direction is no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

4. 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% of the pitch of the respective guide vane or rotor blade cascade.

5. The turbomachine stage as recited in claim 1 wherein the contour varies in the axial direction, the maximum extent of the contour in the axial direction away from the spoke-like pattern being circumferentially spaced from the respective airfoil edge of the respective guide or rotor blade cascade by no more than 25%.

6. The turbomachine stage as recited in claim 5 wherein the maximum variation in the axial direction is no more than 40% of the pitch of the respective guide vane or rotor blade cascade.

7. 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% of the pitch of the respective guide vane or rotor blade cascade.

8. 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.

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

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

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

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

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

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

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

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

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

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

19. The turbomachine stage as recited in claim 1 wherein the phase shift PS(r) varies linearly with the radial position r according to PS(x)=Φ x ×r.

20. The turbomachine stage as recited in claim 1 wherein the phase shift PS(x) varies with the axial position x according to PS(x)=Φ R ×x.

Assignees

• Mtu Aero Engines Gmbh

Inventors

• Mahle Inga
• Gier Jochen
• Koerber Kai
• Engel Karl

Classifications

• F01D11/08

  for sealing space between rotor blade tips and stator (specially-shaped blade tips therefor F01D5/20) · CPC title

• F01D5/141 Primary

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

• F04D29/547

  having a special shape in order to influence fluid flow · CPC title

• F05D2250/70

  Shape · CPC title

• F04D29/321

  for axial flow compressors · CPC title