Mistuned fan for gas turbine engine

Patent metadata

Field	Value
Publication number	US-10837459-B2
Application number	US-201715726819-A
Country	US
Kind code	B2
Filing date	Oct 6, 2017
Priority date	Oct 6, 2017
Publication date	Nov 17, 2020
Grant date	Nov 17, 2020

Abstract

A fan for a gas turbine engine comprises blades distributed around a hub. The blades include first and second blades, having geometric parameters and/or material properties that differ from each other to frequency mistune the fan. The blades are distributed about the hub with at least one second blade between adjacent first blades. The leading edge of the airfoil of the second blades is disposed axially aft of the corresponding leading edge of the first blades in at least a portion of the blade span.

First claim

The invention claimed is:

1. A fan for a gas turbine engine, the fan configured for rotation about a rotational axis defining an axial direction, the fan comprising blades circumferentially distributed around a hub, each of the blades having a leading edge and a trailing edge, the blades including at least a plurality of first blades and a plurality of second blades, the first blades having different geometric parameters and/or material properties than the second blades so as to frequency mistune the first and second blades, and such that the first blades are structurally stronger and more resistant to foreign object damage than the second blades, the first and second blades circumferentially distributed around the hub such that at least one second blades is circumferentially disposed between two adjacent first blades, wherein the leading edge of the at least one second blades is disposed axially aft of the corresponding leading edges of the two adjacent first blades in at least a portion of an outermost 10% of a full span of the respective first and second blades.

2. The fan of claim 1 , wherein the respective first and second blades are defined by a plurality of airfoil sections disposed serially from the hub to tips of the respective first and second blades, the plurality of airfoil sections of the first blades having a distribution of maximum thicknesses from the hub to the tips that is different than a distribution of maximum thicknesses from the hub to the tips of the second blades.

3. The fan of claim 1 , wherein a stiffness of the first blades is greater than a stiffness of the second blades.

4. The fan of claim 1 , wherein a hub-to-tip variation of a chord length of the first blades is different than that of the second blades.

5. The fan of claim 1 , wherein a thickness of the first blades and a thickness of the second blades are different in at least the portion of the outermost 10% of the full span of the respective first and second blades.

6. The fan of claim 1 , wherein a thickness of the first blades at the hub is greater than a thickness of the second blades at the hub.

7. The fan of claim 1 , wherein the leading edge of the at least one second blades is disposed axially aft of the corresponding leading edges of the two adjacent first blades over an entirety of the outermost 10% of the full span, from 90% of the full span to tips of the respective first and second blades.

8. The fan of claim 1 , wherein the leading edge of the at least one second blades is disposed axially aft of the corresponding leading edges of the two adjacent first blades in at least a portion of an outermost 5% of the full span of the respective first and second blades.

9. The fan of claim 8 , wherein the leading edge of the at least one second blades is disposed axially aft of the corresponding leading edges of the two adjacent first blades over an entirety of the outermost 5% of the full span, from 95% of the full span to tips of the respective first and second blades.

10. The fan of claim 1 , wherein the respective first and second blades are circumferentially distributed around the hub in an alternating manner.

11. The fan of claim 1 , wherein a drag coefficient of the first blades is greater than that of the second blades.

12. The fan of claim 1 , wherein the second blades are more aerodynamically efficient than the first blades.

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

14. The fan of claim 1 , wherein the fan is an integrally bladed rotor.

15. The fan of claim 1 , wherein the leading edge of the at least one second blades disposed axially aft of the corresponding leading edges of the two adjacent first blades defines an axial offset, the axial offset corresponding to about 5% to 30% of a chord length of the respective first and second blades.

16. The fan of claim 15 , wherein the axial offset corresponds to about 6% to 8% of the chord length of the respective first and second blades.

17. The fan of claim 1 , wherein the first and second blades each have a respective center of gravity between their respective leading and trailing edges, the center of gravity of the second blades disposed axially aft of the center of gravity of the first blades.

18. A gas turbine engine having a fan configured for rotation about a rotational axis defining an axial direction, the fan comprising blades circumferentially distributed around a hub, each of the blades having a leading edge and a trailing edge, the blades including a first subset of blades and a second subset of blades, the blades of the first subset of blades having different geometric parameters and/or material properties than the second subset of blades, and such that the first subset of blades is structurally stronger and more resistant to foreign object damage than the second subset of blades, the first subset of blades having a natural vibration frequency different than a natural vibration frequency of the second subset of blades so as to frequency mistune the respective blades of the first and second subsets of blades, the blades of the second subset of blades being circumferentially interspaced between blades of the first subset of blades, wherein the leading edges of the blades of the second subset of blades are disposed axially aft of the corresponding leading edges of the blades of the first subset of blades in at least a portion of an outermost 10% of a full span of the respective blades of the first and second subsets of blades.

19. The gas turbine engine of claim 18 , wherein the different geometric parameters and/or material properties of the respective blades may include at least one of a mass, an elastic modulus, a stiffness, a constituent material, a blade thickness, a thickness distribution along the span, a tip profile, and a leading edge profile.

20. The gas turbine engine of claim 18 , wherein the respective blades of the first and second subsets of blades are defined by a plurality of airfoil sections disposed serially from the hub to tips of the respective blades of the first and second subsets of blades, the plurality of airfoil sections of the blades of the first subset of blades having a distribution of maximum thicknesses from the hub to the tips that is different than a distribution of maximum thicknesses from the hub to the tips of the blades of the second subset of blades.

21. The gas turbine engine of claim 18 , wherein a stiffness of the blades of the first subset of blades is greater than a stiffness of the blades of the second subset of blades.

22. The gas turbine engine of claim 18 , wherein a hub-to-tip variation of a chord length of the blades of the first subset of blades is different than that of the blades of the second subset of blades.

23. The gas turbine engine of claim 18 , wherein a thickness of the blades of the first subset of blades and a thickness of the blades of the second subset of blades are different in at least the portion of the outermost 10% of the full span of the respective blades of the first and second subsets of blades.

24. The gas turbine engine of claim 18 , wherein a thickness of the blades of the first subset of blades at the hub is greater than a thickness of the blades of the second subset of blades at the hub.

25. The gas turbine engine of claim 18 , wherein the leading edges of the blades of the second subset of blades are dispo

Assignees

• Pratt & Whitney Canada

Inventors

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

Classifications

• F05D2260/961

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

• F05D2240/301

  Cross-sectional characteristics · CPC title

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title

• F01D5/141

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

• F04D29/324

  Blades · CPC title