U.S. patent application number 13/658209 was filed with the patent office on 2014-04-24 for composite blade with uni-tape airfoil spars.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. The applicant listed for this patent is General Electric Company. Invention is credited to Tod Winton Davis, Nicholas Joseph Kray, Ian Francis Prentice, Pranav Dhoj Shah, Dong-Jin Shim.
Application Number | 20140112796 13/658209 |
Document ID | / |
Family ID | 49326854 |
Filed Date | 2014-04-24 |
United States Patent
Application |
20140112796 |
Kind Code |
A1 |
Kray; Nicholas Joseph ; et
al. |
April 24, 2014 |
COMPOSITE BLADE WITH UNI-TAPE AIRFOIL SPARS
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.
Inventors: |
Kray; Nicholas Joseph;
(Mason, OH) ; Prentice; Ian Francis; (Cincinnati,
OH) ; Davis; Tod Winton; (Liberty Township, OH)
; Shim; Dong-Jin; (Cohoes, NY) ; Shah; Pranav
Dhoj; (Albany, NY) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company; |
|
|
US |
|
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
49326854 |
Appl. No.: |
13/658209 |
Filed: |
October 23, 2012 |
Current U.S.
Class: |
416/226 |
Current CPC
Class: |
F05D 2220/36 20130101;
F05D 2300/603 20130101; F05D 2240/306 20130101; F04D 29/023
20130101; F04D 29/324 20130101; Y02T 50/672 20130101; Y02T 50/60
20130101; F01D 5/16 20130101; F05D 2240/305 20130101; F05D
2300/6034 20130101; F01D 5/282 20130101; F05D 2260/941 20130101;
F01D 5/147 20130101 |
Class at
Publication: |
416/226 |
International
Class: |
F01D 5/28 20060101
F01D005/28 |
Claims
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.
Description
BACKGROUND OF THE INVENTION
[0001] 1. Field of the Invention
[0002] The invention relates to gas turbine engine blades and,
particularly, to composite blades.
[0003] 2. Description of Related Art
[0004] 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.
[0005] 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.
[0006] 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.
[0007] 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
[0008] 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.
[0009] 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.
[0010] 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
[0011] The foregoing aspects and other features of the invention
are explained in the following description, taken in connection
with the accompanying drawings where:
[0012] FIG. 1 is a perspective view illustration of an aircraft gas
turbine engine composite fan blade having a composite uni-tape
spar.
[0013] FIG. 2 is a cross-sectional illustration of the composite
fan blade through 2-2 in FIG. 1.
[0014] FIG. 3 is a perspective diagrammatical view illustration of
an alternative aircraft gas turbine engine composite fan blade
having a composite uni-tape spar.
[0015] FIG. 4 is a perspective diagrammatical view illustration of
the composite uni-tape spar illustrated in FIG. 3.
[0016] 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.
[0017] FIG. 6 is a perspective view illustration of an alternative
aircraft gas turbine engine composite fan blade having a composite
uni-tape spar.
[0018] FIG. 7 is a cross-sectional illustration of the composite
fan blade through 7-7 in FIG. 6.
DETAILED DESCRIPTION OF THE INVENTION
[0019] 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.
[0020] 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.
[0021] 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.
[0022] 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.
[0023] 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 1/3 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.
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] 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.
[0030] 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.
[0031] 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:
* * * * *