Frangible gas turbine engine airfoil with chord reduction

Patent metadata

Field	Value
Publication number	US-10837286-B2
Application number	US-201816161211-A
Country	US
Kind code	B2
Filing date	Oct 16, 2018
Priority date	Oct 16, 2018
Publication date	Nov 17, 2020
Grant date	Nov 17, 2020

Abstract

An airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge. The airfoil includes a blade extending between the root and tip and extending between the leading edge and trailing edge. The blade includes a pressure side and a suction side defining a thickness therebetween at each point along the span. The blade defines a first chord reduction on at least one of the leading edge or trailing edge along a first portion of the span. Further, the blade defines a frangible line extending from the first chord reduction at least partially along the chord at a point along the span within the first portion of the span.

First claim

What is claimed is:

1. An airfoil defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge, the airfoil comprising:
a blade extending between the root and tip and extending between the leading edge and trailing edge, the blade comprising a pressure side and a suction side defining a thickness therebetween at each point along the span, wherein the blade defines a first chord reduction on at least one of the leading edge or trailing edge along a first portion of the span, and wherein the blade defines a frangible line extending from the first chord reduction at least partially along the chord at a point along the span within the first portion of the span;
wherein the first chord reduction is defined on the leading edge, and wherein the blade further defines a second chord reduction on the trailing edge along a second portion of the span, and wherein the frangible line extends between the first chord reduction and the second chord reduction at least partially along the chord at a point along the span within at least one of the first portion or the second portion of the span.

2. The airfoil of claim 1 , wherein the airfoil defines a frangible airfoil portion extending between the frangible line and the tip along the span.

3. The airfoil of claim 1 , wherein the first portion of the span is the same as the second portion of the span.

4. The airfoil of claim 1 , wherein the first portion of the span is positioned toward the tip relative to the second portion of the span.

5. The airfoil of claim 1 , wherein the first portion of the span is positioned toward the root relative to the second portion of the span.

6. The airfoil of claim 1 , wherein the first portion of the span extends to at least 5% of the span from the tip.

7. The airfoil of claim 1 , wherein the first portion of the span extends along at least 10% of the span.

8. The airfoil of claim 1 , wherein the first portion of the span extends along at least 15% but less than 50% of the span.

9. The airfoil of claim 1 , wherein the airfoil further comprises:
a cover sheet coupled to the blade such that the cover sheet covers at least the first chord reduction, wherein the cover sheet defines a continuous leading edge at the first portion of the span.

10. The airfoil of claim 9 , wherein the cover sheet comprises a leading edge sheath.

11. The airfoil of claim 9 , wherein the cover sheet comprises a composite ply including fibers oriented at least partially along at least one of the chord or the thickness.

12. The airfoil of claim 1 , wherein the airfoil further comprises:
a cover sheet coupled to the blade such that the cover sheet covers at least the first chord reduction and the second chord reduction, wherein the cover sheet defines a continuous leading edge at the first portion of the span and a continuous trailing edge at the second portion of the span.

13. The airfoil of claim 1 , wherein the airfoil is a fan blade of a gas turbine engine.

14. The airfoil of claim 1 , wherein the airfoil comprises at least one of a metal, metal alloy, or composite.

15. A gas turbine engine defining a central axis, the gas turbine engine comprising:
an engine shaft extending along the central axis;
a compressor attached to the engine shaft and extending radially about the central axis;
a combustor positioned downstream of the compressor to receive a compressed fluid therefrom;
a turbine mounted on the engine shaft downstream of the combustor to provide a rotational force to the compressor; and
a plurality of airfoils operably connected to the engine shaft, each of the plurality of airfoils defining a span extending between a root and a tip and a chord at each point along the span extending between a leading edge and a trailing edge, each airfoil comprising:
a blade extending between the root and tip and extending between the leading edge and trailing edge, the blade comprising a pressure side and a suction side defining a thickness therebetween at each point along the span, wherein the blade defines a first chord reduction on at least one of the leading edge or trailing edge along a first portion of the span, and wherein the blade defines a frangible line extending from the first chord reduction at least partially along the chord at a point along the span within the first portion of the span;
wherein the first chord reduction is defined on the leading edge, and wherein each blade further defines a second chord reduction on the trailing edge along a second portion of the span, and wherein the frangible line extends between the first chord reduction and the second chord reduction at least partially along the chord at a point along the span within at least one of the first portion or the second portion of the span.

16. The gas turbine engine of claim 15 , further comprising a fan section including the plurality of airfoils configured as fan blades.

17. The gas turbine engine of claim 15 , wherein the first portion of the span extends to at least 5% of the span from the tip.

18. The gas turbine engine of claim 15 , wherein the first portion of the span extends along at least 15% but less than 50% of the span.

Assignees

• Gen Electric

Inventors

• Jain Nitesh
• Kray Nicholas Joseph
• Budumuru Kishore

Classifications

• F04D29/388 Primary

  characterised by construction · CPC title

• Y02T50/60

  Efficient propulsion technologies, e.g. for aircraft · CPC title

• F05D2250/11

  triangular · CPC title

• F02K3/06

  with front fan · CPC title

• F05D2250/141

  circular · CPC title