Vane

Patent metadata

Field	Value
Publication number	US-11415008-B2
Application number	US-202017021435-A
Country	US
Kind code	B2
Filing date	Sep 15, 2020
Priority date	Sep 17, 2019
Publication date	Aug 16, 2022
Grant date	Aug 16, 2022

Abstract

There is disclosed a vane for a gas turbine engine, the vane comprising a three-dimensional weave of composite material. The vane is hollow to define a hollow core, wherein the vane has a maximum thickness (T), and a maximum wall thickness (WT) between an inner surface and an outer surface in a direction normal to the outer surface, wherein WT≤0.1T. Also disclosed is a method of manufacturing a composite vane for a gas turbine engine, the method comprising: forming a preform for the vane comprising a three-dimensional weave of composite material; and curing the preform. The preform has an internal hollow region, and the cured vane has a maximum thickness (T), and a maximum wall thickness (WT) between an inner surface and an outer surface in a direction normal to the outer surface, wherein WT≤0.1T.

First claim

We claim:

1. A vane for a gas turbine engine, the vane comprising a three-dimensional weave of composite material, wherein the vane is hollow to define a hollow core, and wherein the vane has a maximum thickness (T), and a maximum wall thickness between an inner surface and an outer surface in a direction normal to the outer surface (W T ), wherein W T ≤0.1T,
wherein the three-dimensional weave comprises:
a plurality of warp fibres which each extend along a first direction which is within 5 degrees of a spanwise direction of the vane between a root and a tip of the vane, wherein the plurality of warp fibres are grouped into layers of coplanar warp fibres;
a plurality of weft fibres which are grouped into layers of coplanar weft fibres, wherein each layer of coplanar weft fibres extends between successive adjacent layers of coplanar warp fibres, wherein each weft fibre of the plurality of weft fibres extends along a second direction orthogonal to the first direction, wherein each weft fibre extends around an entirety of the hollow core, and wherein each weft fibre is orthogonal to the warp fibres around an entirety of an extent of the weft fibre; and
a plurality of z-binder fibres which extend through a thickness of the warp and weft fibres along a third direction orthogonal to both the first and second directions, each of the plurality of z-binder fibers is generally parallel to the plurality of warp fibres, and wherein the ratio of warp to weft fibres is greater than 50:50.

2. The vane as claimed in claim 1 , wherein the hollow core defines a plurality of hollow core regions which are separated by webbing.

3. The vane as claimed in claim 1 , wherein the vane comprises a damping material within the hollow core.

4. The vane as claimed in claim 1 , wherein a metal or carbon fibre composite rod is provided within the hollow core.

5. The vane as claimed in claim 1 , wherein the vane comprises a connector for connection to an oil line to provide oil to the hollow core.

6. A vane arrangement for a gas turbine engine, comprising
an oil line configured to convey oil; and
the vane as claimed in claim 5 , wherein the oil line is coupled to the connector to provide oil to the hollow core.

7. The vane as claimed in claim 1 , wherein the composite material comprises a carbon fibre reinforcement material and/or an organic polymer matrix.

8. The vane as claimed in claim 1 , wherein the vane comprises an inner region and an outer region, wherein at least one weave parameter differs between the inner region and the outer region, wherein the difference is in a length of the fibres.

9. A fibre reinforced composite stator vane ring for a core inlet, a bypass duct or an air intake of a gas turbine engine, comprising a plurality of the vanes as claimed in claim 1 .

10. A gas turbine engine for an aircraft comprising:
an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;
a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and
wherein a core inlet, a bypass duct or an air intake of the gas turbine engine comprises one or more of the vanes as claimed in claim 1 .

11. The gas turbine engine as claimed in claim 10 , the engine core defining an annular flow path for a core gas flow between a core body and a core fairing, the engine core comprising the core inlet between the core body and the core fairing, wherein the stator vane ring is disposed in the core inlet and serves as a structural support for transferring loads between the core body and the core fairing.

12. The gas turbine engine as claimed in claim 11 , further comprising:
a gearbox that receives an input from the core shaft;
wherein the gearbox receives input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft;
wherein the gearbox is an epicyclic gearbox, and wherein an annulus or ring gear of the epicyclic gearbox is structurally supported by coupling to the stator vane ring.

13. The vane as claimed in claim 1 , wherein the vane comprises an inner region and an outer region, wherein at least one weave parameter differs between the inner region and the outer region, wherein the difference is in a ratio of warp to weft fibres.

14. The vane as claimed in claim 1 , wherein a first layer of coplanar warp fibres includes a first, second, third, and fourth warp fibre extending parallel to each other, wherein a second layer of coplanar warp fibres arranged below the first layer of coplanar warp fibres includes a fifth, sixth, seventh, and eighth warp fibre extending parallel to each other, wherein the first, second, third, and fourth warp fibres are aligned with the fifth, sixth, seventh, and eighth warp fibres, respectively, wherein the plurality of z-binder fibres includes at least one z-binder fibre that includes a first portion that extends over the first warp fibre, a second portion that extends under the second warp fibre and over the sixth warp fibre, a third portion that extends under the seventh warp fibre, and a fourth portion that extends over the eighth warp fibre and under the fourth warp fibre.

15. A method of manufacturing a composite vane for a gas turbine engine, the method comprising:
forming a preform for the vane comprising a three-dimensional weave of composite material; and
curing the preform;
wherein the preform has a hollow core, and
wherein the cured vane has a maximum thickness (T), and a maximum wall thickness between an inner surface and an outer surface in a direction normal to the outer surface (W T ), wherein W T ≤0.1T, and
wherein the three-dimensional weave comprises:
a plurality of warp fibres which each extend along a first direction which is within 5 degrees of a spanwise direction of the vane between a root and a tip of the vane, wherein the plurality of warp fibres are grouped into layers of coplanar warp fibres;
a plurality of weft fibres which are grouped into layers of coplanar weft fibres, wherein each layer of coplanar weft fibres extends between successive adjacent layers of coplanar warp fibres, wherein each weft fibre of the plurality of weft fibres extends along a second direction orthogonal to the first direction, wherein each weft fibre extends around an entirety of the hollow core, and wherein each weft fibre is orthogonal to the warp fibres around an entirety of an extent of the weft fibre; and
a plurality of z-binder fibres which extend through a thickness of the warp and weft fibres along a third direction orthogonal to both the first and second directions, each of the plurality of z-binder fibers is generally parallel to the plurality of warp fibres, and wherein the ratio of warp to weft fibres is greater than 50:50.

16. The method of manufacturing a composite vane as claimed in claim 15 , wherein the preform is formed around a core mandrel, and wherein the core mandrel is removed after curing of the preform, so as to form the hollow core.

17. The method of manufacturing a composite vane as claimed in claim 15 , wherein a damping medium is introduced into the hollow core.

18. The method of manufacturing a composite vane as claimed in claim 15 , wherein the preform comprises an inner region and an outer region, wherein at least one weave parameter differs between the inner region and the outer region, wherein the difference is in a weave dimensionality, namely two dimensional or three-dimensional.

19. The method of manufacturing a composite vane as claimed in claim 15 , wherein the

Assignees

• Rolls Royce Plc

Inventors

• Backhouse Robert C
• Jones Christopher D

Classifications

• F05D2240/12

  Fluid guiding means, e.g. vanes · CPC title

• F05D2300/603

  Composites; e.g. fibre-reinforced · CPC title

• F01D25/34

  Turning or inching gear · CPC title

• F01D9/042

  fixing blades to stators (fixing stator-rings in the casing or to each other F01D25/246) · CPC title

• F01D5/282 Primary

  Selecting composite materials, e.g. blades with reinforcing filaments · CPC title