Mechanically retained fan blade cover

Patent metadata

Field	Value
Publication number	US-10280759-B2
Application number	US-201615163933-A
Country	US
Kind code	B2
Filing date	May 25, 2016
Priority date	May 25, 2016
Publication date	May 7, 2019
Grant date	May 7, 2019

Abstract

A blade assembly may be configured to rotate about an axis of a gas turbine engine. The blade assembly may include a blade and a blade cover. A blade may include a first surface defining a perimeter of a cavity in the blade. A first portion of the perimeter of the cavity includes a lip. A blade cover disposed over the cavity in the blade, such that the cavity is disposed between the blade and the blade cover, and under the lip, such that blade cover is between a portion of the blade and the cavity.

First claim

What is claimed is:

1. A blade assembly configured to rotate about an axis of a gas turbine engine, comprising:
a blade including a first surface defining a perimeter of a cavity in the blade, wherein a first portion of the perimeter of the cavity includes a lip and is disposed along a first leading edge portion, a first chord portion, and a first trailing edge portion, and wherein a second portion of the perimeter of the cavity includes a support surface and is disposed along a second leading edge portion, a second chord portion and a second trailing edge portion, wherein the lip ends at a transition from the first portion of the perimeter of the cavity to the second portion of the perimeter the cavity, and
a blade cover disposed over the cavity in the blade, such that the cavity is disposed between the blade and the blade cover, and under the, such that the blade cover is between a first portion of the blade and the cavity.

2. The blade assembly of claim 1 , wherein the blade cover is disposed over the supporting surface such that a second portion of the blade is disposed between the blade cover and the cavity.

3. The blade assembly of claim 1 , wherein an outer surface of the blade cover contacts an inner surface of the blade, and an inner surface of the blade cover at least one of directly or indirectly contacts an outer surface of the blade.

4. The blade assembly of claim 3 , wherein the inner surface of the blade cover contacts the outer surface of the blade through an adhesive disposed between the blade cover and the blade.

5. A fan section of a gas turbine engine, comprising: a blade assembly configured to rotate about an axis, the blade assembly, comprising:
a blade including a first surface defining a perimeter of a cavity in the blade, wherein a first portion of the perimeter of the cavity includes a lip and is disposed along a first leading edge portion, a first chord portion, and a first trailing edge portion, and wherein a second portion of the perimeter of the cavity includes a support surface and is disposed along a second leading edge portion, a second chord portion and a second trailing edge portion, and
a blade cover disposed over the cavity in the blade, such that the cavity is disposed between the blade and the blade cover, and under the lip, such that the blade cover is between a first portion of the blade and the cavity,
wherein the lip ends at a transition from the first portion of the perimeter of the cavity to the second portion of the perimeter the cavity.

6. The fan section of claim 5 , wherein the blade cover is disposed over the supporting surface such that a second portion of the blade is disposed between the blade cover and the cavity.

7. The fan section of claim 5 , wherein the first chord portion comprises a tip of the blade, and the second chord portion comprises a radially inner edge.

8. The fan section of claim 5 , wherein an outer surface of the blade cover contacts an inner surface of the blade, and an inner surface of the blade cover at least one of indirectly or directly contacts an outer surface of the blade.

9. A gas turbine engine, comprising:
a fan section including:
a disk; and
a blade assembly coupled to the disk, the disk configured to rotate about an axis, the blade assembly comprising:
a blade including a first surface defining a perimeter of a cavity in the blade, wherein a first portion of the perimeter of the cavity includes a lip and is disposed along a first leading edge portion, a first chord portion, and a first trailing edge portion, wherein a second portion of the perimeter of the cavity includes a support surface and is disposed along a second leading edge portion, a second chord portion and a second trailing edge portion, and wherein the lip ends at a transition from the first portion of the perimeter of the cavity to the second portion of the perimeter the cavity, and
a blade cover disposed over the cavity in the blade, such that the cavity is disposed between the blade and the blade cover, and under the lip, such that the blade cover is between a first portion of the blade and the cavity.

10. The gas turbine engine of claim 9 , wherein the first chord portion comprises a tip and the second chord comprises a radially inner edge.

11. The gas turbine engine of claim 9 , wherein an outer surface of the blade cover contacts an inner surface of the blade, and an inner surface of the blade cover contacts an outer surface of the blade.

Assignees

• United Technologies Corp

Inventors

• Murdock James R

Classifications

• Y02T50/672

  Cross-Sectional Technologies · mapped topic

• F05D2260/36

  by a form fit connection, e.g. by interlocking · CPC title

• F01D5/147 Primary

  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

• F01D5/16

  for counteracting blade vibration · CPC title

• Y02T50/673

  Cross-Sectional Technologies · mapped topic