BACKGROUND OF THE INVENTION
FIELD OF THE INVENTION
The invention relates to rotary machine and gas turbine engine rotor and stator
airfoils and, particularly, to composite rotors and stator airfoils.
DESCRIPTION OF RELATED ART
Aircraft turbine engines and other typed of rotary machines include a stationary
and rotating airfoils which channel an airflow downstream. As a result, a wake flow may
be generated and channeled downstream where it may impinge against an object
downstream from the airfoils. Wake flow impingement may generate undesirable noise
and/or aeromechanicalloading. Unwanted noise may be generated by either the
upstream rotating airfoil wake impinging on a stator or rotor component downstream
from the rotating airfoil, or the upstream stator airfoil wake impinging on a rotating
airfoil downstream from the stator airfoil.
The generation of such wake flow may result in a loss of engine performance and
engine efficiency. Reduction of the amplitude of the wake flow may reduce the noise and
the aeromechanicalloading generated when the wake impinges against a downstream
object. An airfoil designed to reduce the amplitude and/or coherence of the wake flow,
the noise, and the aeromechanicalloading is disclosed in United States Patent No.
8,083,487, entitled "AIRFOILS FOR USE IN ROTARY MACHINES AND METHOD
FOR FABRICATING SAME", by Trevor Howard Wood et aI., which issued December
27, 2011 and is incorporated herein by reference. The airfoil includes suction and
pressure sides coupled together at a leading edge and a trailing, wherein the airfoil
includes a plurality of first and second chord sections each extending between the trailing
and leading edges, wherein at least one of the first chord sections extends outward from
the pressure side of the airfoil at the trailing edge, and at least one of the second chord
sections extends outward from the suction side of the airfoil at the trailing edge.
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Particular embodiments of the airfoil are wavy or crenelated airfoils.
Composite fan blades have been developed for aircraft gas turbine engines to
reduce weight and cost, particularly for 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. The term composite as used herein may be defined
as a material containing a reinforcement such as fibers or particles supported in a binder
or matrix material. Composites include metallic and non-metallic composites. One
particularly useful embodiment for fan composite fan blades is made of a unidirectional
tape material and an epoxy resin matrix. The composite fan blade and other airfoils
disclosed herein may include composite materials of 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, PMRI5, 8MI, PEEU, etc. A more
particular material includes fibers unidirectionally aligned into 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
within.
It is highly desirable to provide light-weight and strong aircraft gas turbine engine
fan blades that also reduce the amplitude of wake flow, noise, and aeromechanical
loading.
SUMMARY OF THE INVENTION
A gas turbine engine airfoil includes chordwise spaced apart leading and trailing
edges, pressure and suction sides extending outwardly in a spanwise direction from an
airfoil base to an airfoil tip, trailing edge cladding made of a cladding material bonded to
a composite core made of a composite material, the cladding material being less brittle
than the composite material, the composite core including a central core portion
extending chordwise downstream from a leading edge portion to a trailing edge portion
of the composite core, and the trailing edge cladding including a wavy wall and the
trailing edge.
The airfoil may include comprising pressure and suction side flanks of the trailing
edge cladding bonded to pressure and suction side surfaces respectively of the trailing
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edge portion of the composite core. Waves of the wavy wall may extend normal to and
away from the pressure and suction side surfaces. The metallic trailing edge cladding
may include spanwise extending wavy pressure and suction side trailing edge guards
which include the waves of the wavy wall.
An erosion coating may be used to cover the composite core and butt up against
and hide the forward facing steps on the pressure and suction side flanks of the trailing
edge cladding. Alternatively, rebates may extend into the composite core and hide the
steps.
A gas turbine engine fan blade may include the airfoil extending outwardly from a
platform of the blade. The blade may include a root extending inwardly from the
platform and the root may include an integral dovetail.
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 longitudinal part sectional and part diagrammatical view illustration of
an exemplary embodiment of an aircraft turbofan gas turbine engine with a composite
core fan blade having a metallic wavy wall trailing edge.
FIG. 2 is a perspective view illustration of the composite core fan blade illustrated
in FIG. 1.
FIG. 3 is a cross-sectional diagrammatical view illustration of a composite core
trailing edge and metallic trailing edge of the blade taken through 3-3 in FIG. 2.
FIG. 4 is a cross-sectional view illustration of a first alternative embodiment of
the composite core trailing edge and metallic trailing edge ofthe blade illustrated in FIG.
3.
FIG. 5 is a cross-sectional view illustration of a second alternative embodiment of
the composite core trailing edge and metallic trailing edge of the blade illustrated in FIG.
3.
DETAILED DESCRIPTION OF THE INVENTION
Illustrated in FIG. 1 is an exemplary aircraft turbofan gas turbine engine 10
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circumscribed about an engine centerline axis 12 and suitably designed to be mounted to
a wing or fuselage of an aircraft. The engine 10 includes, in downstream serial flow
communication, a fan 14, a booster 16, a high pressure compressor 18, a combustor 20, a
high pressure turbine (HPT) 22, and a low pressure turbine (LPT) 24. The HPT or high
pressure turbine 22 is joined by a high pressure drive shaft 23 to the high pressure
compressor 18. The LPT or low pressure turbine 24 is joined by a low pressure drive
shaft 25 to both the fan 14 and the booster 16.
In typical operation, air 26 is pressurized by a row of fan blades 11 in the fan 14
and produces an inner air flow 15 channeled through the booster 16 which further
pressurizes the inner air flow 15. The pressurized air is then flowed to the high pressure
compressor 18 which further pressurizes the air. The pressurized air is mixed with fuel in
the combustor 20 for generating hot combustion gases 28 that flow downstream in tum
through the HPT 22 and the LPT 24.
A flow splitter 34 surrounding the booster 16 immediately behind the fan 14
includes a sharp leading edge 32 which splits the fan air 26 pressurized by the fan 14 into
a radially inner stream (inner air flow 15) channeled through the booster 16 and a radially
outer stream (bypass air flow 17) channeled through the bypass duct 36. A fan casing 30
surrounding the fan 14 is supported by an annular fan frame 33. The booster 16 includes
alternating annular rows of booster blades and vanes 38, 42 extending radially outwardly
and inwardly across a booster flowpath 39 in a booster duct 40. The annular rows of
booster blades 38 are suitably joined to the fan 14. The booster 16 is located forward of
the fan frame 33 and is disposed radially inboard of the flow splitter 34. The fan 14
includes a plurality of fan blades 11 that extend substantially radially outwardly from a
fan rotor disk 13.
Illustrated in FIG. 2 is one embodiment of the fan blade 11 that may be used in
engine 10 (illustrated in FIG. 1). The fan blade 11 includes an airfoil 45 extending
outwardly from a platform 56 and a root 54 extending inwardly from the platform 56.
Alternatively, airfoil 45 may be used with, but not limited to, rotor blades, stator vanes,
and/or nozzle assemblies. Airfoil 45 may also be used with, OGVs and the booster.
In the exemplary embodiment, the root 54 includes an integral dovetail 58 that
enables the fan blade 11 to be mounted to the rotor disk 13. The airfoil 45 includes
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pressure and suction sides 41, 43 extending outWardly in a spanwise direction along a
span S from an airfoil base 49 at the platform 56 to an airfoil tip 47. The exemplary
pressure and suction sides 41,43 illustrated herein are concave and convex respectively.
The airfoil 45 extends along a chord C between chordwise spaced apart leading and
trailing edges LE, TE. The airfoil 45 may be mounted on and integral with a hub instead
of the platform and disk to form an integrally bladed rotor (IBR). Alternatively, fan
blade 11 may have any conventional form, with or without dovetail 58 or platform 56.
For example, fan blade 11 may be formed integrally with disk 13 in a blisk-type
configuration that does not include the dovetail 58 and the platform is annular extending
around the entire blisk.
Referring to FIGS. 2 and 3, the airfoil 45 includes a composite core 44 and
trailing edge cladding 46 that provides the airfoil's trailing edge TE. Demarcation line 59
indicates the intersection of the composite core 44 and metallic trailing edge cladding 46.
The composite core 44 is made of a composite material, generally airfoil shaped, and
incudes a central core portion 63 extending chordwise downstream from a leading edge
portion 48 to a trailing edge portion 50 of the composite core 44. The trailing edge
cladding 46 is made of any suitable material that is stronger or more ductile or less brittle
than the composite material of the composite core 44.
The trailing edge cladding material is illustrated herein as being metallic.
Another less brittle and suitable cladding material is S-glass such as HS2 and HS4 which
are high strength glass fibers made from magnesium alumina silicate. The leading edge
portion 48 mayor may not be covered by leading edge cladding 66 made of a metallic or
other suitable material and which would then define the leading edge LE ofthe airfoil 45.
The trailing edge cladding 46 includes a fluted or wavy wall 70 and the trailing
edge TE designed to reduce noise during the engine's operation and, thus, lower the
acoustic signature of the airfoil 45. The wavy wall 70 is designed to mix the fan wakes to
reduce the wake interaction with downstream outlet guide vanes (DGV). The wavy wall
70 includes waves 68 such as crenelations or undulations 72. This allows the peak strain
caused by the wavy wall 70, which is an aerodynamic feature stress, to be born by the
metallic trailing edge cladding 46. The metallic trailing edge cladding 46 is far more
capable of bearing strain than the composite core 44.
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Referring to FIG. 3, the metallic trailing edge cladding 46 is bonded to the trailing
edge portion 50 ofthe composite core 44. The bonding uses a film adhesive for example.
The trailing edge cladding 46 includes pressure and suction side flanks 73, 74 that are
bonded to pressure and suction side surfaces 76, 78 respectively of the trailing edge
portion 50 of the composite core 44. The waves 68 of the wavy wall 70 extend normal to
and away from the pressure and suction side surfaces 76, 78. Thus, the wavy wall 70,
representative of a sculptured trailing edge (STE) feature, is entirely made of metal and
better able to withstand peak stresses and strains that peak in the STE feature itself than
composite airfoils or composite portions of airfoils. The metallic trailing edges have a
higher strain capability as compared to the composite core based on pure material
property evaluation. Bonding metallic STE features to the composite core allows stresses
transferred to the composite core to be spread out over a large area and, thus, lowering
localized stress and strain that may fail a composite airfoil. The airfoil 45 with the
composite core 44 and the metallic trailing edge cladding 46 provides a more capable
metallic material which increases robustness of the airfoil and provides weight
advantages of composite materials.
The exemplary embodiment of the metallic trailing edge cladding 46 illustrated in
FIG. 3 includes radially or spanwise extending wavy pressure and suction side trailing
edge guards 80, 82 made of sheet metal. The pressure and suction side trailing edge
guards 80, 82 provide the wavy shape of the metallic trailing edge cladding 46. The
metal cladding may be hot formed to shape. The pressure and suction side trailing edge
guards 80, 82 are bonded to the pressure and suction side surfaces 76, 78 respectively of
the composite core 44. Contacting pressure and suction side portions 86, 88 of the
pressure and suction side trailing edge guards 80, 82 are bonded together as indicated
along bond line 89.
The pressure and suction side trailing edge guards 80, 82 diagrammatically
illustrated in FIG. 3 have an upstream or forward facing step 90 that should be
aerodynamically covered. Thus, this part ofthe metallic cladding is blended into the
composite core 44. FIG. 4 illustrates one design to hide the step 90 by providing a small
rebate 92 into the composite core 44 to keep the pressure and suction side flanks 73, 74
flush with the composite core 44. Alternatively, as diagrammatically illustrated in FIG.
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5, an erosion coating 96 butting up to the step 90 may be placed on the composite core
44.
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:

WE CLAIM:
1. A gas turbine engine airfoil (45) comprising:
chordwise spaced apart leading and trailing edges (LE, TE),
pressure and suction sides (41, 43) extending outwardly in a spanwise direction
from an airfoil base (49) to an airfoil tip (47),
trailing edge cladding (46) made of a cladding material bonded to a composite
core (44) made ofa composite material wherein the cladding material is less brittle than
the composite material,
the composite core (44) including a central core portion (63) extending chordwise
downstream from a leading edge portion (48) to a trailing edge portion (50) of the
composite core (44), and
the trailing edge cladding (46) including a wavy wall (70) and the trailing edge
(TE).
2. The airfoil (45) as claimed in Claim 1, further comprising pressure and suction
side flanks (73, 74) of the trailing edge cladding (46) bonded to pressure and suction side
surfaces (76, 78) respectively of the trailing edge portion (50) of the composite core (44).
3. The airfoil (45) as claimed in Claim 1, further comprising waves (68) of the wavy
wall (70) extending normal to and away from the pressure and suction side surfaces (76,
78).
4. The airfoil (45) as claimed in Claim 1, further comprising the metallic trailing
edge cladding (46) including spanwise extending wavy pressure and suction side trailing
edge guards (80,82) including the waves (68) of the wavy wall (70).
5. The airfoil (45) as claimed in Claim 4, further comprising the wavy pressure and
suction side trailing edge guards (80, 82) being made of sheet metal.
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6. The airfoil (45) as claimed in Claim 5, further comprising pressure and suction
side flanks (73, 74) of the trailing edge cladding (46) bonded to pressure and suction side
surfaces (76, 78) respectively of the trailing edge portion (50) of the composite core (44).
7. The airfoil (45) as claimed in Claim 6, further comprising waves (68) ofthe wavy
wall (70) extending normal to and away from the pressure and suction side surfaces (76,
78).
8. The airfoil (45) as claimed in Claim 6, further comprising forward facing steps
(90) on the pressure and suction side flanks (73, 74) of the trailing edge cladding (46) and
an erosion coating (96) covering the composite core (44) and butting up against and
hiding the steps (90).
9. The airfoil (45) as claimed in Claim 6, further comprising forward facing steps
(90) on the pressure and suction side flanks (73, 74) of the trailing edge cladding (46) and
rebates (92) extending into the composite core (44) and hiding the steps (90).
10. A gas turbine engine fan blade (11) comprising:
an airfoil (45) extending outwardly from a platform (56) of the blade (11),
the airfoil (45) extending between chordwise spaced apart leading and trailing
edges (LE, TE),
the airfoil (45) including pressure and suction sides (41, 43) extending outwardly
in a spanwise direction from an airfoil base (49) at the platform (56) to an airfoil tip (47),
the airfoil (45) including trailing edge cladding (46) made of a cladding material
bonded to a composite core (44) made of a composite material wherein the cladding
material is less brittle than the composite material,
the airfoil (45) including the composite core (44) including a central core portion
(63) extending chordwise downstream from a leading edge portion (48) to a trailing edge
portion (50) of the composite core (44), and
the trailing edge cladding (46) including a wavy wall (70) and the trailing edge
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(TE).
11. The blade (11) as claimed in Claim 10, further comprising a root (54) extending
inwardly from the platform (56).
12. The blade (11) as claimed in Claim 11, further comprising the root (54) including
an integral dovetail (58).