Fan blades for gas turbine engines with reduced stress concentration at leading edge

Patent metadata

Field	Value
Publication number	US-9617860-B2
Application number	US-201213721650-A
Country	US
Kind code	B2
Filing date	Dec 20, 2012
Priority date	Dec 20, 2012
Publication date	Apr 11, 2017
Grant date	Apr 11, 2017

Abstract

A light weight fan blade for a gas turbine engine is disclosed. The disclosed fan blade has an increased chord length at a transition section between the inner edge or single tooth attachment portion of the fan blade that connects to the rotor disk and the airfoil. Increasing the chord length at the transition section substantially reduces the stress concentration when the fan blade is operated under normal operating conditions. Reducing the stress concentration enables engineers to design lighter fan blades and larger fan blades for larger gas turbine engines. By increasing the size of gas turbine engines, additional thrust can be provided without adding additional engines to the aircraft.

First claim

The invention claimed is:

1. A fan blade comprising:
an airfoil connected to a single tooth attachment;
a transition section extending between a first junction defined by the airfoil and a first end of the transition section and a second junction defined by a second end of the transition section disposed opposite the first end of the transition section and the single tooth attachment, the transition section having a concave profile that extends from the second junction towards a transition point and a convex profile that extends from the transition point towards the first junction;
a first chord extending between the first junction and a trailing edge ofthe single tooth attachment; and
a second chord spaced apart from the first chord, the second chord extending between the second junction and the trailing edge of the single tooth attachment, the first chord being longer than the second chord.

2. The fan blade of claim 1 wherein the airfoil is at least substantially covered by a sheath.

3. The fan blade of claim 2 wherein the sheath does not cover the transition section.

4. The fan blade of claim 2 wherein the sheath is fabricated from at least one of titanium and a titanium alloy.

5. The fan blade of claim 2 , wherein the sheath defines a flow path section extending along an arcuate path from the first junction towards a trailing edge of the airfoil.

6. A gas turbine engine comprising:
a fan blade assembly including a rotor disk having an outer periphery, the outer periphery of the rotor disk including a plurality of single tooth attachment slots;
a plurality of fan blades, each fan blade including an airfoil connected to a single tooth attachment, each single tooth attachment slot of the rotor disk accommodating the single tooth attachment of one of the fan blades;
for each fan blade, the airfoil, the single tooth attachment, and a transition section extending between the airfoil and the single tooth attachment including leading and trailing edges;
for each fan blade, the transition section defining a first chord at a junction between the transition section and the airfoil, for each fan blade, the transition section defining a second chord at a junction between the transition section and the single tooth attachment;
wherein for each fan blade, the transition section extends laterally outward away from the single tooth attachment from the junction between the transition section and the single tooth attachment before reaching the junction between the transition section and the airfoil, the transition section having a concave profile and a convex profile that extends from the concave profile, and a discontinuity between the transition section and a leading edge of the fan blade.

7. The engine of claim 6 wherein, for each fan blade, the trailing edge of its airfoil, transition section and single tooth attachment are at least substantially coplanar.

8. The engine of claim 6 wherein each airfoil of each fan blade is at least substantially covered by a sheath.

9. The engine of claim 8 wherein each sheath does not cover the transition section of its fan blade.

10. The engine of claim 8 wherein each sheath is fabricated from a material selected from the group consisting of titanium and a titanium alloy.

11. The engine of claim 6 wherein, for each fan blade, the leading edge of the transition section curves as said leading edge extends from the junction of said transition section and single tooth attachment of said fan blade to the junction of said transition section and airfoil of said fan blade.

12. The engine of claim 6 wherein, for each fan blade, the junction of the transition section and the airfoil has a first thickness and the junction of the transition section and the single tooth attachment has a second thickness, and
wherein for each fan blade, the second thickness is greater than the first thickness.

Assignees

• United Technologies Corp

Inventors

• Lattanzio Santiago
• Amoo Olaleye M

Classifications

• Y02T50/671

  Cross-Sectional Technologies · mapped topic

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title

• F01D5/147

  Construction, i.e. structural features, e.g. of weight-saving hollow blades (F01D5/148, F01D5/16 and F01D5/20 take precedence; blade shape F01D5/141; blades with cooling or heating channels or cavities F01D5/18; heating, heat-insulating or cooling means on blades F01D5/18) · CPC title

• Y02T50/673

  Cross-Sectional Technologies · mapped topic

• F01D5/3007 Primary

  of axial insertion type · CPC title