BACKGROUND OF THE INVENTION
This invention relates generally to composite components and more
particularly to the configuration of mounting features of composite components such
as turbomachinery airfoils.
It is desirable to manufacture gas turbine components such as turbomachinery
blades from composite materials that provide favorable strength-to-weight ratios.
Known types of composite materials include polymer matrix composites ("PMC"),
typically suitable for fan blades, and ceramic matrix composites ("CMC"), typically
suitable for turbine blades.
All of these composite materials are comprised of a laminate of a matrix
material and reinforcing fibers and are orthotropic to at least some degree, i.e. the
material's tensile strength in the direction parallel to the length of the fibers (the "fiber
direction") is stronger than the tensile strength in the perpendicular direction (the
"matrix" or "interlarninar" direction). The physical properties such as modulus and
Poisson's ratio also differ between the fiber and matrix. The primary fiber direction in
turbomachinery blades is typically aligned with the radial or spanwise direction in
order to provide the greatest strength capability to carry the centripetal load imparted
by the spinning rotor. As such, the weaker matrix, secondary or tertiary (i.e. nonprimary)
fiber direction is then orthogonal to the radial direction.
As composites have different coefficients of thermal expansion ("CTE") than
metal alloys use for the rotor disk, all of the blade dovetails use a configuration that
allows for free thermal expansion between the two parts. However, this type of
dovetail configuration leads to a peak interlaminar tensile stress imparted in the shank
of the composite blade, which must be carried in the weaker matrix material, just
above the pressure faces of the dovetail, commonly referred to as the "minimum
neck", which can be the limiting stress location in the blade design.
The matrix, or non-primary fiber direction strength, herein referred to as
interlaminar strength, is typically weaker (i.e. 1/10 or less) than the fiber direction
strength of a composite material system and can be the limiting design feature on
composite blades, in particular, CMC turbine blades.
Accordingly, there is a need for a blade mounting structure which reduces
interlaminar stresses in the mounting attachment region for a composite blade.
BRIEF DESCRIPTION OF THE INVENTION
This need is addressed by the present invention, which provides a
turbomachinery blade structure that includes first and second minimum necks
configured to produce reduced interlaminar tensile stresses during operation.
According to an aspect of the invention a turbomachinery blade includes: an
airfoil; and a shank extending from a root of the airfoil, the shank being constructed
from a composite material including reinforcing fibers embedded in a matrix, wherein
the shank includes a pair of spaced-apart side faces. The side faces cooperatively
define: a dovetail disposed at a radially inboard end of the shank, comprising spacedapart,
diverging faces; a first neck portion having a concave curvature disposed
radially outboard of the dovetail, and defining a primary minimum neck at which a
thickness of the shank is at a local minimum; and a second neck portion disposed
radially outboard of the first minimum neck, the second neck portion having a
concave curvature and defining a secondary minimum neck at which the thickness of
the shank is at a local minimum.
BRIEF DESCRIPTION OF THE DRAWINGS
The invention may be best understood by reference to the following
description taken in conjunction with the accompanying drawing figures in which:
FIG. 1 is a perspective view of a turbine blade of a gas turbine engine;
FIG. 2 is a schematic, transverse sectional view of a shank portion of a prior
art turbine blade; and
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FIG. 3 is a schematic, transverse sectional view of a shank portion of a turbine
blade constructed according to an aspect of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
Referring to the drawings wherein identical reference numerals denote the
same elements throughout the various views, FIG. 1 illustrates an exemplary lowpressure
turbine (or "LPT") blade 22. While illustrated and explained in the context of
a LPT blade, it will be understood that the principles of the present invention are
equally applicable to other types of turbomachinery airfoils, such as fan and
compressor blades, high-pressure turbine ("HPT") blades, or stationary airfoils.
The turbine blade 22 is constructed from a composite material such as a CMC
or PMC material, described in more detail below. The turbine blade 22 includes a
dovetail 36 configured to engage a dovetail slot 38 (see FIG. 3) of a gas turbine
engine rotor disk 24 of a known type, for radially retaining the turbine blade 22 to the
rotor disk 24 as it rotates during operation. The dovetail 36 is an integral part of a
blade shank 40. The shape of the shank 40 transitions from the dovetail 36 to the
curved airfoil shape to allow for a smooth transition for composite layup. A platform
42 projects laterally outwardly from and surrounds the shank 40. The platform 42 may
be integral to the turbine blade 22 or may be a separate component. An airfoil 44
extends radially outwardly from the shank 40. The airfoil 44 has a concave pressure
side 46 and a convex suction side 48 joined together at a leading edge 50 and at a
trailing edge 52. The airfoil 44 has a root 54 and a tip 56, which may incorporate a tip
shroud. The airfoil 44 may take any configuration suitable for extracting energy from
the hot gas stream and causing rotation of the rotor disk.
For comparison purposes, FIG. 2 shows a schematic view of a shank 140 of a
prior art turbine blade. The shank 140 includes spaced-apart generally parallel left and
right side faces 158. At the radially inner end (or inboard end), The side faces 158
define a dovetail 136 having a pair of spaced-apart, divergent pressure faces 160. A
concave-curved transition section 166 is disposed just outboard of the dovetail 136.
The portion of the shank 140 where the transition section 166 meets the remainder of
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the side faces 158 constitutes a "minimum neck" 164. The thickness of the shank 140
in the tangential direction "T" is at a minimum at the location of the minimum neck
164. In operation, the primary load on the rotating turbine blade is in the radial (or
spanwise) direction "R". As a result of blade radial force, the turbine blade is also
subject to tensile stresses in the tangential direction T, caused by interaction of the
pressure faces 160 with the dovetail slot 138 of a turbine rotor disk 124. The
tangential stresses are of a much lower magnitude than the spanwise stresses. for
example, the maximum radial, fiber, stresses may be about 10 times greater than the
maximum tangential stresses. In a prior art turbine blade constructed from a isotropic,
or near isotropic (i.e. directionally solidified) metal alloy, this does not present a
problem as strengths in any direction are equivalent.
However, as noted above, composite materials are typcially orthotropic to at
least some degree. For example, the yield strength or the ultimate tensile strength of a
composite material could exhibit a 10: 1 or 15: 1 ratio between the radial (fiber) and
tangential (matrix or interlaminar) directions.
Accordingly, the shank 40 of the turbine blade 22 seen in FIGS. 1 and 3 is
configured to reduce the interlaminar stresses in the composite material that forms the
turbine blade 22. FIG. 3 shows a schematic view of a portion of the shank 40.
The shank 40 includes spaced-apart left and right side faces 58 which are
contoured in a specific manner, and may be described as having several distinct
"portions". At the radially inner end (or inboard end), The side faces 58 define the
dovetail 36 that includes a pair of spaced-apart, divergent pressure faces 60.
Just outboard of the dovetail 36, there is a first neck portion 62. In the first
neck portion 62, each side face 58 defines a concave curve. At the radially outer end
of the first neck portion 62, it defines a first (or primary) minimum neck 64, where the
thickness of the shank 40 in the tangential direction T is at a local minimum relative
to the immediately surrounding structure. As used herein the term "minimum neck"
does not necessarily imply any specific dimensions. The portions of the side faces 58
defining the first or primary minimum neck 64 have a first radius "Rl".
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Just outboard (or radially outward) of the primary minimum neck 64, there is a
first transition portion 66. In the first transition portion 66, each side face 58 defines a
smooth convex curve. Other configurations of the side faces 58 which could produce
similar results include straight lines or spline shapes.
Outboard of first transition portion 66, there is a second or secondary neck
portion 68. In the secondary neck portion 68, each side face 58 defines a smooth
concave curve having a second radius "R2". The radius R2 is larger than the radius
R1. The secondary neck portion 68 defines a second (or secondary) minimum neck
70, where the thickness of the shank 40 in the tangential direction T is at a local
minimum relative to the immediately surrounding structure.
A second transition portion 72 is disposed outboard of the secondary neck
portion 68. In the second transition portion 72, each side face 58 defines a smooth
convex curve. Other configurations of the side faces 58 which could produce similar
results include straight lines or spline shapes.
An outboard portion 74 is disposed outboard of the second transition portion.
In the outboard portion 74, the side faces 58 are generally parallel to each other as
they transition to the airfoil geometry.
The profile of the side faces 58 is shaped so as to be compatible with
composite materials. The reinforcing fibers generally follow the contours of (i.e. are
parallel to) the side faces 58. The side faces 58 are contoured such that the fibers will
not buckle or wrinkle where outward cusps are located. While the profile of the side
faces 58 has been illustrated as exemplary two-dimensional sectional views, it is
noted that the actual shape may be different at each axial section. In other words,
applicability to actual 3D blade shanks will follow this configuration described above,
but adds another dimension to tailor the geometry.
In the illustrated example, the thickness of the shank 40 in the tangential
direction "T" is significantly less (from a functional standpoint) at the location of the
secondary minimum neck 70 than at the primary minimum neck 64. The exact shapes
and dimensions of the side faces 58 may be altered to suit a particular application and
the specific composite material used.
Generally, PMC materials are highly orthotropic. One example of a known
PMC is a carbon fiber reinforced epoxy, which would typically be used in a fan blade.
Other fiber materials such as boron or silicon carbide are also known. Other matrix
materials such as phenolic, polyester, and polyurethane for example, are known as
well.
Generally, CMC materials are less orthotropic than PMC materials, and may
be have properties which are close to isotropic. Examples of known CMC materials
include a ceramic type fiber for example Sic, forms of which are coated with a
compliant material such as Boron Nitride (BN). The fibers are carried in a ceramic
type matrix, one form of which is Silicon Carbide (Sic). CMC materials would
typically be suitable for a turbine blade.
By addition of a secondary minimum neck 70 above the primary minimum
neck 64 the shank interlaminar stiffness is softened to allow the resultant interlaminar
stress to be distributed over a larger area, thus reducing the peak interlaminar tensile
stress value. Analysis has shown that the shank configuration described above can
lower the peak interlaminar tensile stress by a significant amount, for example about
20% to 30%, as compared to the prior art configuration. This configuration can be
used to add design margin at the minimum neck of the blade in order to enable
designs to be able carry more radial loads, via larger engine radius or higher speed
applications, or to add interlaminar stress margin to existing blade designs.
This configuration also enables additional high cycle fatigue ("HCF")
capability for blades by allowing the vibratory modes of the blade which have
inflection at or near the primary minimum neck per the prior art sketch (i.e. 1st flex or
IF), to then inflect about the thinner net section of the secondary minimum neck,
which has a lower radial static stress due to the larger radius and associated lower
stress concentration factor, to enable a larger allowance for HCF stress.
The foregoing has described an interlaminar stress reducing configuration
for composite turbine components. While specific embodiments of the present
invention have been described, it will be apparent to those skilled in the art that
various modifications thereto can be made without departing from the spirit and scope
of the invention. Accordingly, the foregoing description of the preferred embodiment
of the invention and the best mode for practicing the invention are provided for the
purpose of illustration only and not for the purpose of limitation.

WE CLAIM:
1. A turbomachinery blade, comprising:
an airfoil; and
a shank extending from a root of the airfoil, the shank being constructed from
a composite material including reinforcing fibers embedded in a matrix, wherein the
shank includes a pair of spaced-apart side faces, the side faces cooperatively defining:
a dovetail disposed at a radially inboard end of the shank, comprising
spaced-apart, diverging faces;
a first neck portion having a concave curvature disposed radially
outboard of the dovetail, and defining a primary minimum neck at which a thickness
of the shank is at a local minimum; and
^jtjjf a second neck portion disposed radially outboard of the first minimum
neck, the second neck portion having a concave curvature and defining a secondary
minimum neck at which the thickness of the shank is at a local minimum.
2. The turbomachinery blade of claim 1 wherein:
the first neck portion has a first radius; and
the second neck portion has a second radius substantially greater than the first
radius.
3. The turbomachinery blade of claim 1 wherein the thickness of the shank at
the second neck portion is significantly less than the thickness at the first neck
portion.
4. The turbomachinery blade of claim 1 wherein the airfoil includes: leading
and trailing edges extending between a root and a tip, and opposed pressure and
suction sides joined together at the leading and trailing edges.
5. The turbomachinery blade of claim 1 wherein a first transition portion is
disposed between the first neck portion and the second neck portion, and wherein the
side faces are convex-curved within the first transition portion.
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6. The turbomachinery blade of claim 5 wherein a second transition portion is
disposed outboard of the second neck portion, and wherein the side faces are convexcurved
within the first transition portion.
7. The turbomachinery blade of claim 6 wherein an outboard portion is
disposed outboard of the second transition portion, and wherein the side faces are
generally parallel to each other within the outboard portion.
8. The tiu-bomachinery blade of claim 1 wherein the composite material has a
strength ratio of fiber direction to matrix direction of at least about 10 to 1.
9. The turbomachinery blade of claim 1 wherein the composite material is a
polymer matrix composite.
10. The turbomachinery blade of claim 1 wherein the composite material is a
ceramic matrix composite.