Partially coated blade

Patent metadata

Field	Value
Publication number	US-10174626-B2
Application number	US-201414514697-A
Country	US
Kind code	B2
Filing date	Oct 15, 2014
Priority date	Oct 15, 2014
Publication date	Jan 8, 2019
Grant date	Jan 8, 2019

Abstract

A partially coated blade for a gas turbine engine, including a fillet surface surrounding the airfoil section and connecting it to the platform section. A radially outermost portion of the pressure side and leading edge is covered by a thermal barrier coating. This portion extends radially from a first limit to the blade tip. The first limit is located at a radial distance from the platform of at most 21% of the maximum span. The fillet surface is free or substantially free of the thermal barrier coating. In another embodiment, a second portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, extending radially from the platform section to a second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span. A method of applying a thermal barrier coating is also discussed.

First claim

The invention claimed is:

1. A partially coated blade for a gas turbine engine, the blade comprising:
a platform section;
an airfoil section extending radially outwardly from the platform section and defining a tip of the blade,
a maximum radial distance between the platform section and the tip defining a maximum span of the blade, the airfoil section including:
opposed leading and trailing edges,
a pressure side extending from the platform section to the tip and interconnecting the leading and trailing edges, and
a suction side from the platform section to the tip and interconnecting the leading and trailing edges opposite the pressure side;

wherein a first portion of the pressure side and of the leading edge is covered by a thermal barrier coating, the first portion extending radially between a first limit and the tip, the first limit located at a radial distance from the platform section of at most 21% of the maximum span;
wherein a second portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, the second portion extending radially between the platform section and a second limit, the second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span; and
wherein the suction side includes a third portion extending axially from the trailing edge toward the leading edge, the third portion being free or substantially free of the thermal barrier coating, an upstream transition zone extending axially on the suction side between the third portion and the first portion, the upstream transition zone including a tapering thickness of the thermal barrier coating to define a transition between the third portion and the first portion, the upstream transition zone extending axially from the leading edge along a distance of from 14% to 18% of a maximum chord length of the airfoil.

2. The blade as defined in claim 1 , wherein the second limit is located a radial distance from the platform section corresponding to at least 14% of the maximum span.

3. The blade as defined in claim 1 , wherein an inner transition zone is defined between the first and second limits, the inner transition zone including a tapering thickness of the thermal barrier coating to define a transition between the first and second portions.

4. The blade as defined in claim 3 , wherein:
the trailing edge is free or substantially free of the thermal barrier coating; and
a downstream transition zone extends axially on the pressure side between the first portion and the trailing edge, the downstream transition zone extending radially between the tip and the inner transition zone, the downstream transition zone including a tapering thickness of the thermal barrier coating to define a transition between the first portion and the trailing edge.

5. The blade as defined in claim 4 , wherein the downstream transition zone extends axially along a distance of from 9% to 11% of a maximum chord length of the airfoil section.

6. The blade as defined in claim 3 , wherein:
the upstream transition zone extends radially between the tip and the inner transition zone.

7. The blade as defined in claim 1 , wherein the second portion includes a thickness of the thermal barrier coating of 0.001 inch or less.

8. The blade as defined in claim 7 , wherein the first portion includes a thickness of the thermal barrier coating of from 0.003 inch to 0.007 inch.

9. A partially coated blade for a gas turbine engine, the blade comprising:
a platform section;
an airfoil section extending radially outwardly from the platform section and defining a tip of the blade, the airfoil section having:
opposed leading and trailing edges,
a pressure side interconnecting the leading and trailing edges, and
a suction side interconnecting the leading and trailing edges opposite the pressure side;

a rounded fillet surface surrounding the airfoil section and connecting the leading edge, trailing edge, pressure side and suction side to the platform section, the fillet surface extending radially from an outer end tangential to the airfoil section to an inner end tangential to the platform section;
wherein a radially outermost portion of the pressure side and of the leading edge is covered by a thermal barrier coating, the radially outermost portion extending radially between a first limit and the tip, the first limit located at a radial distance from the platform section of at most 21% of the maximum span;
wherein the fillet surface is free or substantially free of the thermal barrier coating; and
wherein the suction side includes a portion extending axially from the trailing edge toward the leading edge, the portion of the suction side being free or substantially free of the thermal barrier coating, an upstream transition zone extending axially on the suction side between the portion of the suction side and the radially outermost portion, the upstream transition zone including a tapering thickness of the thermal barrier coating to define a transition between the portion of the suction side and the radially outermost portion, the upstream transition zone extending axially from the leading edge along a distance of from 14% to 18% of a maximum chord length of the airfoil.

10. The blade as defined in claim 9 , wherein a radially innermost portion of the pressure side and of the leading edge is free or substantially free of the thermal barrier coating, the radially innermost portion extending radially between the platform section and a second limit, the second limit located a radial distance from the platform section corresponding to at least 5% of the maximum span.

11. The blade as defined in claim 10 , wherein the second limit is located a radial distance from the platform section corresponding to at least 14% of the maximum span.

12. The blade as defined in claim 10 , wherein an inner transition zone is defined between the first and second limits, the inner transition zone including a tapering thickness of the thermal barrier coating to define a transition between the radially innermost and radially outermost portions.

13. The blade as defined in claim 12 , wherein:
the trailing edge is free or substantially free of the thermal barrier coating; and
a downstream transition zone extends axially on the pressure side between the radially outermost portion and the trailing edge, the downstream transition zone extending radially between the tip and the inner transition zone, the downstream transition zone including a tapering thickness of the thermal barrier coating to define a transition between the radially outermost portion and the trailing edge.

14. The blade as defined in claim 13 , wherein:
the upstream transition zone extending radially between the tip and the inner transition zone.

15. The blade as defined in claim 9 , wherein a radially innermost portion of the leading edge and of the pressure side includes a thickness of the thermal barrier coating of 0.001 inch or less.

16. The blade as defined in claim 15 , wherein the radially outermost portion includes a thickness of the thermal barrier coating of from 0.003 inch to 0.007 inch.

17. A method of applying a thermal barrier coating on a blade, the method comprising:
masking at least a filler surface defining a connection between a platform section and an airfoil section of the blade, the airfoil section having a pressure side and a suction side extending chord-wise between a leading edge and a trailing edge; and
applying the thermal barrier coating on a first portion of the pressure side an

Assignees

• Pratt & Whitney Canada

Inventors

• Papple Michael
• Lecuyer Daniel

Classifications

• Y02T50/60

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

• F05D2230/90

  Coating; Surface treatment · CPC title

• F01D5/288 Primary

  Protective coatings for blades · CPC title

• F05D2240/303

  related to the leading edge of a rotor blade · CPC title

• F05D2240/305

  related to the pressure side of a rotor blade · CPC title