Blade made of composite material with variable-density attached leading edge

Patent metadata

Field	Value
Publication number	US-12037924-B2
Application number	US-202017784932-A
Country	US
Kind code	B2
Filing date	Dec 14, 2020
Priority date	Dec 18, 2019
Publication date	Jul 16, 2024
Grant date	Jul 16, 2024

Abstract

A method for manufacturing a blade from composite material having an added metal leading edge for a gas turbine aeroengine includes producing a blade body from composite material including a blade root part, a shank part and an airfoil body part; manufacturing, via additive manufacturing, a leading edge extending in a longitudinal direction between a lower end present at the shank part of the composite material and an upper end present at the tip; bonding the manufactured leading edge onto the foremost edge portion of the airfoil body of the blade body.

First claim

The invention claimed is:

1. A method for manufacturing a blade in composite material having an added metal leading edge for a gas turbine aeroengine, the method comprising:
producing a blade body from composite material comprising a blade root part, a shank part and an airfoil body part, the airfoil body part extending in a longitudinal direction between the shank part and an airfoil body tip and in a transverse direction between a foremost edge portion and a rear edge portion,
additively manufacturing the metal leading edge extending in the longitudinal direction between a lower end to be present at the shank part of the blade body in composite material and an upper end to be present at the tip of the airfoil body part, and
bonding the metal leading edge onto the foremost edge portion of the airfoil body,
wherein the metal leading edge comprises a first portion in a first metal material and extending from the lower end of the leading edge as far as an intermediate position located between the lower end and upper end of the leading edge, and a second portion in a second metal material extending from the intermediate position as far as the upper end of the leading edge in an upper part of the blade, the second metal material having a greater density than a density of the first metal material so that the leading edge has a greatest stiffness in the upper part of the blade.

2. The method according to claim 1 , wherein the first portion extends in the longitudinal direction over a height corresponding to 30% of a total airflow height while the second portion extends in the longitudinal direction over a height corresponding to 70% of the total airflow height.

3. The method according to claim 1 , wherein the first metal material is steel or titanium, while the second metal material is a nickel or cobalt alloy.

4. The method according to claim 1 , wherein the first metal material is titanium while the second metal material is steel.

5. The method according to claim 1 , wherein the first portion extends in the longitudinal direction over a first height and the second portion extends in the longitudinal direction over a second height that is greater than the first height.

6. A blade in composite material having an added metal leading edge for a gas turbine aeroengine, the blade comprising a blade structure in composite material comprising a blade root, a shank and an airfoil body, the airfoil body extending in a longitudinal direction between the shank and a blade tip and in a transverse direction between a foremost edge portion and a rear edge portion, and a metal leading edge bonded onto the foremost edge portion of the airfoil body of the blade structure in composite material, the leading edge extending in the longitudinal direction between a lower end present at the shank of the blade structure in composite material and an upper end present at the blade tip of said blade structure, wherein the metal leading edge comprises a first portion in a first metal material and extending from the lower end of the leading edge as far as an intermediate position located between the lower end and upper end of the leading edge, and a second portion in a second metal material and extending from the intermediate position as far as the upper end of said leading edge, the second metal material having a greater density than a density of the first metal material so that the leading edge has a greatest stiffness in the upper part of the blade.

7. The blade according to claim 6 , wherein the first portion extends in the longitudinal direction over a height corresponding to 30% of a total airflow height while the second portion extends in the longitudinal direction over a height corresponding to 70% of the total airflow height.

Assignees

• Safran Aircraft Engines

Inventors

• Gondre Guillaume Pascal Jean-Charles
• Bazot Olivier
• Dubosc Matthieu Pierre Michel

Classifications

• B29C66/7394

  characterised by the material of at least one of the parts being a thermoset · CPC title

• B29C66/7392

  characterised by the material of at least one of the parts being a thermoplastic · CPC title

• B29C66/742

  to metals or their alloys · CPC title

• B29C66/721

  Fibre-reinforced materials (B29C66/729 takes precedence) · CPC title

• B29C66/53

  Joining single elements to tubular articles, hollow articles or bars · CPC title