Rotor with non-uniform blade tip clearance

Patent metadata

Field	Value
Publication number	US-10408231-B2
Application number	US-201715703472-A
Country	US
Kind code	B2
Filing date	Sep 13, 2017
Priority date	Sep 13, 2017
Publication date	Sep 10, 2019
Grant date	Sep 10, 2019

Abstract

A rotor for a gas turbine engine comprises a rotor having a hub and blades around the hub, and extending from the hub to tips. The tips include first and second tip portions between their respective tip leading edge and tip trailing edge. Tips are spaced from a rotational axis of the rotor by spans. A mean span of a first tip portion of a first blade is greater than a mean span of a corresponding first tip portion of a second blade. A mean span of a second tip portion the first blade is less than a mean span of a corresponding second tip portion of the second blade.

First claim

The invention claimed is:

1. A rotor for a gas turbine engine, the rotor adapted to be received within a casing having a radially inner surface and configured for rotation about a rotational axis, the rotor comprising a hub and blades circumferentially distributed around the hub, the blades extending radially along spans from the hub to tips thereof and including at least first blades and second blades, the blades having airfoils with leading edges and trailing edges, the tips of the blades extending axially relative to the rotational axis of the rotor from tip leading edges to tip trailing edges, the tips of each of the blades having at least first and second tip portions extending axially between the tip leading edges and the tip trailing edges; wherein a mean span of the first tip portion of the first blades is less than a mean span of the corresponding first tip portion of the second blades, and a mean span of the second tip portion of the first blades is greater than a mean span of the corresponding second tip portion of the second blades.

2. The rotor of claim 1 , wherein the spans vary from the tip leading edges to the tip trailing edges.

3. The rotor of claim 1 , wherein the span of the first blades increases from the tip leading edge to the tip trailing edge thereof, and the span of the second blades decreases from the tip leading edge to the tip trailing edge thereof.

4. The rotor of claim 1 , wherein each of the first blades is disposed circumferentially between two of the second blades, the first blades having a natural vibration frequency different than a natural vibration frequency of the second blades.

5. The rotor of claim 1 , wherein the first tip portions extend downstream from the tip leading edges and the second tip portions extend upstream from the tip trailing edges.

6. The rotor of claim 1 , wherein a ratio of a maximum span difference between spans of the first blades and of the second blades over a mean diameter of the rotor is from 0.0001 to 0.001.

7. The rotor of claim 1 , wherein the first blades have a natural vibration frequency different than a natural vibration frequency of the second blades.

8. The rotor of claim 1 , wherein the first and second tip portions meet between the tip leading and trailing edges.

9. A gas turbine engine comprising: a rotor having a hub and a plurality of blades circumferentially distributed around the hub, the blades extending radially from the hub to tips of the blades, the blades having airfoils with leading edges and trailing edges, the tips of the blades extending axially relative to a rotational axis of the rotor from tip leading edges to tip trailing edges, the tips of the blades having at least first and second tip portions extending between the tip leading edges and the tip trailing edges; and a casing disposed around the rotor, a radially-inner surface of the casing spaced from the tips of the blades by radial tip clearances; wherein a mean radial tip clearance of a first tip portion of one of the blades is greater than a mean radial tip clearance of a first tip portion of another one of the blades, and a mean radial tip clearance of a second tip portion the one of the blades is less than a mean radial tip clearance of a second tip portion of the other one of the blades.

10. The gas turbine engine of claim 9 , wherein radial tip clearances of the blade tips vary from the tip leading edges to the tip trailing edges.

11. The gas turbine engine of claim 9 , wherein a radial tip clearance of the one of the blades decreases from a tip leading edge to a tip trailing edge thereof and a radial tip clearance of the other one of the blades increases from a tip leading edge to a tip trailing edge thereof.

12. The gas turbine engine of claim 9 , wherein the blades include first blades and second blades, each of the first blades disposed circumferentially between two of the second blades, the first blades having a natural vibration frequency different than a natural vibration frequency of the second blades, the one of the blades being one of the first blades, the other one of the blades being one of the second blades.

13. The gas turbine engine of claim 9 , wherein the first tip portions extend downstream from the tip leading edges and the second tip portions extend upstream from the tip trailing edges.

14. The gas turbine engine of claim 9 , wherein a ratio of a maximum radial tip clearance difference between radial tip clearances of the one of the blades and of the other one of the blades over a diameter of the rotor is from 0.001 to 0.0001.

15. The gas turbine engine of claim 9 , wherein the one of the blades has a natural vibration frequency different than a natural vibration frequency of the other one of the blades.

16. The gas turbine engine of claim 9 , wherein the first and second tip portions meet between the tip leading and trailing edges.

17. A method of forming a rotor within a casing of a gas turbine engine, the method comprising:
providing the rotor with a hub and a plurality of blades circumferentially distributed around the hub, the blades extending radially from the hub to tips of the blades and including at least first and second blades, the tips of the blades adapted to be circumscribed by the casing;
forming a first radial tip clearance gap between a first tip portion of the first blades and a layer of abradable material on an inner surface of the casing; and
forming a second radial tip clearance gap between a second tip portion of the second blades and the layer of abradable material, the first and second radial tip clearance gaps being different.

18. The method of claim 17 , wherein a mean radial tip clearance of first tip portions of the first blades is greater than a mean radial tip clearance of first tip portions of the second blades, and a mean radial tip clearance of second tip portions of the first blades is less than a mean radial tip clearance of second tip portions of the second blades.

19. The method of claim 17 , wherein the first blades have a natural vibration frequency different than a natural vibration frequency of the second blades.

20. The method of claim 17 , wherein, during operation, the first blades axially deflect relative to the second blades.

Assignees

• Pratt & Whitney Canada

Inventors

• Veitch Thomas
• Abrari Farid
• Adique Ernest
• Fudge Daniel
• Heikurinen Kari
• Stone Paul
• Theratil Ignatius
• Townsend Peter
• Urac Tibor

Classifications

• F05D2240/307

  related to the tip of a rotor blade · CPC title

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title

• F05D2260/961

  by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape · CPC title

• F04D29/324

  Blades · CPC title

• F01D5/16

  for counteracting blade vibration · CPC title