US20140112796A1

US20140112796A1 - Composite blade with uni-tape airfoil spars

Info

Publication number: US20140112796A1
Application number: US13/658,209
Authority: US
Inventors: Nicholas Joseph Kray; Ian Francis Prentice; Tod Winton DAVIS; Dong-Jin Shim; Pranav Dhoj Shah
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2012-10-23
Filing date: 2012-10-23
Publication date: 2014-04-24
Also published as: CA2888777A1; EP2917496A1; JP6179961B2; WO2014065968A1; JP2015537143A; CN104981586A; BR112015009060A2

Abstract

A gas turbine engine composite blade includes an airfoil having pressure and suction sides extending outwardly in a spanwise direction from a blade root along a span to a blade tip. A core section of the blade including composite quasi-isotropic plies extends spanwise outwardly through the blade. One or more spars including a stack of uni-tape plies having predominately a 0 degree fiber orientation with respect to the span and extending spanwise outwardly through the root and a portion of the airfoil towards the tip. Spars may include pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil and which be located near or along the pressure and suction sides respectively. Chordwise extending portion may be centered about a maximum thickness location of the airfoil. Spars may have height, width, and thickness that avoids flexural airfoil modes.

Description

BACKGROUND OF THE INVENTION

1. Field of the Invention
The invention relates to gas turbine engine blades and, particularly, to composite blades.
2. Description of Related Art
Composite blades made from elongated filaments composited in a light-weight matrix have been developed for aircraft gas turbine engines. The blades are light-weight having high strength. The term composite has come to be defined as a material containing a reinforcement such as fibers or particles supported in a binder or matrix material. Many composites are used in the aerospace industry including both metallic and non-metallic composites. The composites used for the blades disclosed herein are made of a unidirectional tape material and an epoxy resin matrix. A discussion of this and other suitable materials may be found in the “Engineering Materials Handbook” by ASM INTERNATIONAL, 1987-1989 or later editions.
The composite blades disclosed herein are made from the non-metallic type made of a material containing a fiber such as a carbonaceous, silica, metal, metal oxide, or ceramic fiber embedded in a resin material such as Epoxy, PMR15, BMI, PEEU, etc. The fibers are unidirectionally aligned in a tape that is impregnated with a resin, formed into a part shape, and cured via an autoclaving process or press molding to form a light weight, stiff, relatively homogeneous article having laminates or plies within.
Composite fan blades have been developed for aircraft gas turbine engines to reduce weight and cost, particularly, for fan blades in larger engines. A large engine composite wide chord fan blades offer a significant weight savings over a large engine having standard chorded fan blades. Among the problems, all gas turbine engine blades face resonance or flexural modes. Large composite fan blades for high bypass ratio aircraft gas turbine engines with relatively wide diameter fans are faced with this problem. This is particularly true for the frequencies that cause the blade to experience first and second flexural airfoil modes.
It is highly desirable to provide light-weight and strong aircraft gas turbine engine fan blades that avoid passing through or experiencing assonance and flexural modes and, in particular, first and second flexural airfoil modes.

SUMMARY OF THE INVENTION

A gas turbine engine composite fan blade includes an airfoil having pressure and suction sides extending outwardly in a spanwise direction from a blade root of the blade along a span to a blade tip. A core section of the blade includes composite quasi-isotropic plies extending spanwise outwardly through the blade including the root and the airfoil towards the tip. One or more spars including a stack of uni-tape plies having a preferential 0 degree fiber orientation with respect to the span spanwise outwardly through the root and through a portion of the airfoil towards the tip.
The chordwise extending portion of the core section may be centered about a maximum thickness location of the airfoil. The spars may have a spanwise height, chordwise width, and spar thickness that avoids flexural airfoil modes such as first and second flexural airfoil modes. The one or more spars may include pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil which may be located near or along the pressure and suction sides respectively.
In one embodiment of the blade, the one or more spars include chordwise spaced apart upstream and downstream pressure side spars and chordwise spaced apart upstream and downstream suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

BRIEF DESCRIPTION OF THE DRAWINGS

The foregoing aspects and other features of the invention are explained in the following description, taken in connection with the accompanying drawings where:

FIG. 1 is a perspective view illustration of an aircraft gas turbine engine composite fan blade having a composite uni-tape spar.

FIG. 2 is a cross-sectional illustration of the composite fan blade through 2-2 in FIG. 1.

FIG. 3 is a perspective diagrammatical view illustration of an alternative aircraft gas turbine engine composite fan blade having a composite uni-tape spar.

FIG. 4 is a perspective diagrammatical view illustration of the composite uni-tape spar illustrated in FIG. 3.

FIG. 5 is perspective diagrammatical view illustration of −P degree, 0 degree, and +P degree plies of the composite fan blade illustrated in FIG. 2.

FIG. 6 is a perspective view illustration of an alternative aircraft gas turbine engine composite fan blade having a composite uni-tape spar.

FIG. 7 is a cross-sectional illustration of the composite fan blade through 7-7 in FIG. 6.

DETAILED DESCRIPTION OF THE INVENTION

Illustrated in FIGS. 1 and 2 is a composite fan blade 10 for a high bypass ratio fanjet gas turbine engine (not shown) having a composite airfoil 12. Composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of filament reinforced composite plies 40 (illustrated in FIG. 5). As used herein, the terms “lamination” and “ply” are synonymous. The airfoil 12 includes pressure and suction sides 41, 43 extending outwardly in a spanwise direction from a fan blade root 20 along a span S to a blade tip 47. In the exemplary embodiment, the root 20 includes an integral dovetail 28 that enables the fan blade 10 to be mounted to a rotor disk.
The exemplary pressure and suction sides 41, 43 illustrated herein are concave and convex respectively. The airfoil 12 extends along a chord C between chordwise spaced apart leading and trailing edges LE, TE. Thickness T of the airfoil 12 varies in both chordwise and spanwise directions C, S and extends between pressure and suction sides 41, 43 of the blade 10 also referred to as convex and concave sides of the blade or airfoil. The airfoil 12 may be mounted on and be integral with a hub to form an integrally bladed rotor (IBR) or integrally with a disk in a BLISK configuration.
The plies 40 are generally all made from a unidirectional fiber filament ply material, preferably a tape, as it is often referred to, arranged generally in order of span and used to form a composite airfoil 12 as shown in FIG. 1. The plies 40 are essentially those plies that form the airfoil 12 and root 20 of the blade 10 as illustrated in FIGS. 1 and 3.
The composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of different filament reinforced airfoil plies 40. The blade 10 uses filament reinforced laminations or plies with a filament orientation of 0 degrees, +P degrees, and −P degrees as illustrated in FIG. 5. The angle P is a predetermined angle as measured from 0 degrees which corresponds to a generally radially extending axis of the airfoil which may be its centerline or stacking line and is typically about 45 degrees. An exemplary arrangement is more particularly pointed out and explained in U.S. Pat. No. 4,022,547 by Stanley.
Referring to FIGS. 1-4, the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies 52. Pressure and suction side spars 54, 56 sandwich a chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies 52 generally near or along the pressure and suction sides 41, 43 respectively in the airfoil 12. The chordwise extending portion 58 of the core section 50 extends chordwise partially through the airfoil 12. The chordwise extending portion 58 of the core section 50 is generally centered chordwise in the airfoil 12. The exemplary embodiment of the chordwise extending portion 58 illustrated herein extends chordwise about ⅓ through the airfoil 12 and is generally centered chordwise about in the middle of the airfoil 12. The composite quasi-isotropic ply chordwise extending portion 58 of the core section 50 is preferably limited to a thicker cross sectional area of the airfoil 12 around or centered about a maximum thickness Tmax location 61 of the airfoil 12, as illustrated in FIG. 2, so as to be most effective. The Tmax location 61 is about a middle third of the airfoil between the leading and trailing edges LE, TE in the chordwise direction C for the exemplary airfoil illustrated herein. The pressure and suction side spars 54, 56 are made from stacks 62 of preferential 0 degree uni-tape plies 63 (see FIG. 5) with a 0 degree fiber orientation with respect to the span S.
Referring to FIGS. 3 and 4, the pressure and suction side spars 54, 56 (and the uni-tape plies they are made from extend) spanwise S through the fan blade root 20 and through a portion 53 of the airfoil 12 to a spar tip 57. The pressure and suction side spars 54, 56 have a spanwise height H as measured from the fan blade root 20 to the spar tip 57 which is less than the span S of the airfoil. In the embodiment of the composite fan blade 10 illustrated herein, the pressure and suction side spars 54, 56 (and the uni-tape plies are made from) extend all the way through the root 20 including the dovetail 28.
The quasi-isotropic ply core section 50 generally include alternating plies of tape with different +P, 0, and −P fiber orientations. The pressure and suction side spars 54, 56 include uni-tape plies with a predominately 0 degree fiber orientation. An exemplary blade ply lay-up is disclosed in U.S. Pat. No. 5,375,978, entitled “Foreign Object Damage Resistant Composite Blade and Manufacture” to Evans, which issued Dec. 27, 1994, is assigned to the same assignee of this patent, and is incorporated herein by reference. The ply lay-up disclosed in U.S. Pat. No. 5,375,978 is referred to as a standard quasi-isotropic lay-up sequence of 0. degree, +45 degree, 0 degree, −45 degree fiber orientations with the plies having the numerous ply shapes.
The stacks 62 of the spars include uni-tape plies with a predominately 0 degree fiber orientation. A few of the plies may have another fiber orientation. An example is a stack having a total of 8 plies with 4 plies of 0 degree fiber orientation on both sides of two plies having +30 and a −30 degree plies. This ply layup may be represented or denoted by 0,0,0,0,+30,−30,0,0,0,0.
Referring to FIGS. 1, 2, and 3, the spars have a spanwise height H, chordwise width W, and spar thickness TS designed to increase radial or spanwise stiffness of the airfoil 12 without increasing the weight of the blade. The spars are also designed or tailored or tuned to avoid flexural airfoil modes such as first and second flexural airfoil modes 1F and 2F. The spanwise height H and the spar thickness TS are designed or tailored to tuned or avoid flexural airfoil modes such as first and second flexural airfoil modes 1F and 2F. The uni-tape ply spar with predominately a 0 degree fiber orientation allows for a stiffer blade without adding thickness and without adding weight and performance penalties. The exemplary embodiment of the composite blade illustrated herein is a fan blade but the composite blade with a quasi-isotropic ply core section and spars made from stacks 62 of 0 degree uni-tape plies 63 may also be used for other gas turbine engine blades such as compressor blades.
The exemplary embodiment of the composite blade 10 illustrated herein includes one or more outer cover plies 66 around the core section 50, made of composite quasi-isotropic plies, and the pressure and suction side spars 54, 56. A leading edge metallic shield 68 is bonded around the leading edge LE. The shield is often referred to as metallic cladding.
Referring to FIGS. 6 and 7, an alternative spar design for the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies and two sets of pressure and suction side spars. The two sets include chordwise spaced apart upstream and downstream pressure side spars 74, 76 and chordwise spaced apart upstream and downstream suction side spars 78, 80 sandwiching the chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies generally near or along the pressure and suction sides 41, 43 respectively.
The present invention has been described in an illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation. While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims:

Claims

What is claimed is:

1. A gas turbine engine composite blade comprising:

an airfoil having pressure and suction sides extending outwardly in a spanwise direction from a blade root of the blade along a span to a blade tip,

a core section of the blade including composite quasi-isotropic plies extending spanwise outwardly through the blade including the root and the airfoil towards the tip,

one or more spars including a stack of uni-tape plies having predominately a 0 degree fiber orientation with respect to the span, and

the one or more spars extending spanwise outwardly through the root and through a portion of the airfoil towards the tip.

2. The blade as claimed in claim 1 further comprising the one or more spars including pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

3. The blade as claimed in claim 2 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

4. The blade as claimed in claim 1 further comprising the chordwise extending portion of the core section centered about a maximum thickness location of the airfoil.

5. The blade as claimed in claim 4 further comprising the one or more spars including pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

6. The blade as claimed in claim 5 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

7. The blade as claimed in claim 1 further comprising the spars having a spanwise height, chordwise width, and spar thickness that avoids flexural airfoil modes.

8. The blade as claimed in claim 7 further comprising the flexural airfoil modes including first and second flexural airfoil modes.

9. The blade as claimed in claim 8 further comprising the one or more spars including pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

10. The blade as claimed in claim 9 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

11. The blade as claimed in claim 8 further comprising the chordwise extending portion of the core section centered about a maximum thickness location of the airfoil.

12. The blade as claimed in claim 11 further comprising the one or more spars including pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

13. The blade as claimed in claim 12 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

14. The blade as claimed in claim 1 further comprising the one or more spars including chordwise spaced apart upstream and downstream pressure side spars and chordwise spaced apart upstream and downstream suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

15. The blade as claimed in claim 14 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

16. The blade as claimed in claim 15 further comprising the chordwise extending portion of the core section centered about a maximum thickness location of the airfoil.

17. The blade as claimed in claim 16 further comprising the upstream and downstream pressure side spars and the chordwise spaced apart upstream and downstream suction side spars having a spanwise height, chordwise width, and spar thickness that avoids flexural airfoil modes.

18. The blade as claimed in claim 17 further comprising the flexural airfoil modes including first and second flexural airfoil modes.

19. The blade as claimed in claim 17 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.

20. The blade as claimed in claim 19 further comprising the flexural airfoil modes including first and second flexural airfoil modes.

21. The blade as claimed in claim 1 further comprising:

the root includes an integral dovetail,

one or more outer cover plies around the core section, and

a leading edge metallic shield bonded around the leading edge.

22. The blade as claimed in claim 21 further comprising the spars having a spanwise height, chordwise width, and spar thickness that avoids flexural airfoil modes.

23. The blade as claimed in claim 22 further comprising the flexural airfoil modes including first and second flexural airfoil modes.

24. The blade as claimed in claim 23 further comprising the one or more spars including pressure and suction side spars sandwiching a chordwise extending portion of the core section in the airfoil.

25. The blade as claimed in claim 24 further comprising the chordwise extending portion of the core section centered about a maximum thickness location of the airfoil.

26. The blade as claimed in claim 25 further comprising the chordwise extending portion of the core section located near or along the pressure and suction sides respectively.