[0001] The disclosure relates generally to turbomachines and, more particularly, to
a cast turbine nozzle having heat transfer protrusions on an inner surface of a
leading edge of a cooling cavity in the airfoil.
BACKGROUND
[0002] Turbine nozzles include cooling cavities in airfoil bodies to direct a coolant
to cool the airfoil. The cooling cavity provides space for an impingement cooling
sleeve that directs coolant against an inner surface of the airfoil body that defines
the cooling cavity. In certain nozzle stages, it is advantageous to make leading
edges of the turbine nozzles smaller in radius, which narrows the airfoil. The
narrower airfoils make it more difficult to maintain cooling with conventional
impingement cooling.
BRIEF DESCRIPTION
[0003] A first aspect of the disclosure provides a cast turbine nozzle including: an
airfoil having a body including a suction side, a pressure side opposing the suction
side, a leading edge spanning between the pressure side and the suction side, a
trailing edge opposing the leading edge and spanning between the pressure side and
the suction side, and a cooling cavity defined by an inner surface of the body; at
least one endwall connected with the airfoil along the suction side, the pressure side,
the trailing edge and the leading edge; and a plurality of heat transfer protrusions
extending inwardly from the inner surface of the body within the cooling cavity,
the plurality of heat transfer protrusions extending from the leading edge along the
suction side and along the pressure side in a radially staggered columnar pattern,
wherein the inner surface includes a planar surface extending between adjacent heat
transfer protrusions.
[0004] A second aspect of the disclosure provides a nozzle section for a turbine, the
nozzle section having a set of nozzles, the set of nozzles including at least one cast
nozzle having: an airfoil having a body including a suction side, a pressure side
opposing the suction side, a leading edge spanning between the pressure side and
3
the suction side, a trailing edge opposing the leading edge and spanning between
the pressure side and the suction side, and a cooling cavity defined by an inner
surface of the body; at least one endwall connected with the airfoil along the suction
side, the pressure side, the trailing edge and the leading edge; and a plurality of heat
transfer protrusions extending inwardly from the inner surface of the body within
the cooling cavity, the plurality of heat transfer protrusions extending from the
leading edge along the suction side and along the pressure side in a radially
staggered columnar pattern, wherein the inner surface includes a planar surface
extending between adjacent heat transfer protrusions.
[0005] A third aspect of the disclosure provides a turbine including a plurality of
cast turbine nozzles, each of the cast turbine nozzles comprising: an airfoil having
a body including a suction side, a pressure side opposing the suction side, a leading
edge spanning between the pressure side and the suction side, a trailing edge
opposing the leading edge and spanning between the pressure side and the suction
side, and a cooling cavity defined by an inner surface of the body; at least one
endwall connected with the airfoil along the suction side, the pressure side, the
trailing edge and the leading edge; and a plurality of heat transfer protrusions
extending inwardly from the inner surface of the body within the cooling cavity,
the plurality of heat transfer protrusions extending from the leading edge along the
suction side and along the pressure side in a radially staggered columnar pattern,
wherein the inner surface includes a planar surface extending between adjacent heat
transfer protrusions.
[0006] The illustrative aspects of the present disclosure are designed to solve the
problems herein described and/or other problems not discussed.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] These and other features of this disclosure will be more readily understood
from the following detailed description of the various aspects of the disclosure taken
in conjunction with the accompanying drawings that depict various embodiments
of the disclosure, in which:
[0008] FIG. 1 is a schematic view of an illustrative turbomachine in the form of a
4
combustion turbine or gas turbine (GT) system, according to embodiments of the
disclosure;
[0009] FIG. 2 is a cross-section illustration of an example gas turbine assembly
with a four-stage turbine that may be used with the turbomachine in FIG. 1;
[0010] FIG. 3 shows a schematic perspective view of an illustrative pair of turbine
nozzles including an airfoil with heat transfer projections, according to various
embodiments of the disclosure;
[0011] FIG. 4 shows a perspective view of an illustrative impingement sleeve for
use with the turbine nozzle, according to embodiments of the disclosure;
[0012] FIG. 5 shows an overhead perspective view of a pair of cast turbine nozzles
in a turbine nozzle section, according to embodiments of the disclosure;
[0013] FIG. 6 shows a slightly enlarged overhead perspective view of a cast turbine
nozzle, according to embodiments of the disclosure;
[0014] FIG. 7 shows a perspective view of a number of heat transfer projections,
according to embodiments of the disclosure;
[0015] FIG. 8 shows a plan view of an inner surface of a cooling cavity looking at
the top of heat transfer projections, according to embodiments of the disclosure;
and
[0016] FIG. 9 shows a cross-sectional side view of heat transfer projections along
line 9-9 in FIG. 8, according to embodiments of the disclosure.
[0017] It is noted that the drawings of the disclosure are not necessarily to scale.
The drawings are intended to depict only typical aspects of the disclosure and
therefore should not be considered as limiting the scope of the disclosure. In the
drawings, like numbering represents like elements between the drawings.
DETAILED DESCRIPTION
[0018] As an initial matter, in order to clearly describe the subject matter of the
current disclosure it will become necessary to select certain terminology when
referring to and describing relevant machine components within a turbomachine.
To the extent possible, common industry terminology will be used and employed
in a manner consistent with its accepted meaning. Unless otherwise stated, such
5
terminology should be given a broad interpretation consistent with the context of
the present application and the scope of the appended claims. Those of ordinary
skill in the art will appreciate that often a particular component may be referred to
using several different or overlapping terms. What may be described herein as
being a single part may include and be referenced in another context as consisting
of multiple components. Alternatively, what may be described herein as including
multiple components may be referred to elsewhere as a single part.
[0019] In addition, several descriptive terms may be used regularly herein, and it
should prove helpful to define these terms at the onset of this section. These terms
and their definitions, unless stated otherwise, are as follows. As used herein,
“downstream” and “upstream” are terms that indicate a direction relative to the flow
of a fluid, such as coolant in a post-impingement space in an airfoil or, for example,
the flow of air through the combustor. The term “downstream” corresponds to the
direction of flow of the fluid, and the term “upstream” refers to the direction
opposite to the flow. The terms “forward” and “aft,” without any further specificity,
refer to directions, with “forward” referring to the front or compressor end of the
engine, and “aft” referring to the rearward section of the turbomachine.
[0020] It is often required to describe parts that are disposed at differing radial
positions with regard to a center axis. The term “radial” refers to movement or
position perpendicular to an axis. For example, if a first component resides closer
to the axis than a second component, it will be stated herein that the first component
is “radially inward” or “inboard” of the second component. If, on the other hand,
the first component resides further from the axis than the second component, it may
be stated herein that the first component is “radially outward” or “outboard” of the
second component. The term “axial” refers to movement or position parallel to an
axis. Finally, the term “circumferential” refers to movement or position around an
axis. It will be appreciated that such terms may be applied in relation to the center
axis of the turbine.
[0021] In addition, several descriptive terms may be used regularly herein, as
described below. The terms “first”, “second”, and “third” may be used
interchangeably to distinguish one component from another and are not intended to
6
signify location or importance of the individual components.
[0022] The terminology used herein is for the purpose of describing particular
embodiments only and is not intended to be limiting of the disclosure. As used
herein, the singular forms “a”, “an” and “the” are intended to include the plural
forms as well, unless the context clearly indicates otherwise. It will be further
understood that the terms “comprises” and/or “comprising,” when used in this
specification, specify the presence of stated features, integers, steps, operations,
elements, and/or components but do not preclude the presence or addition of one or
more other features, integers, steps, operations, elements, components, and/or
groups thereof. “Optional” or “optionally” means that the subsequently described
event or circumstance may or may not occur or that the subsequently described
element or feature may or may not be present and that the description includes
instances where the event occurs (or the feature is present) and instances where it
does not (or is not present).
[0023] Where an element or layer is referred to as being “on,” “engaged to,”
“connected to” or “coupled to” another element or layer, it may be directly on,
engaged to, connected to, or coupled to the other element or layer, or intervening
elements or layers may be present. In contrast, when an element is referred to as
being “directly on,” “directly engaged to,” “directly connected to” or “directly
coupled to” another element or layer, there may be no intervening elements or layers
present. Other words used to describe the relationship between elements should be
interpreted in a like fashion (e.g., “between” versus “directly between,” “adjacent”
versus “directly adjacent,” etc.). As used herein, the term “and/or” includes any
and all combinations of one or more of the associated listed items.
[0024] Embodiments of the disclosure provide a cast turbine nozzle, a turbine
nozzle section, and a turbine. The turbine nozzle includes a plurality of heat transfer
protrusions on an inner surface of a cooling cavity in the airfoil thereof. The heat
transfer protrusions provide improved cooling effectiveness to maintain part life,
turbine efficiency, and power output. More particularly, the heat transfer
protrusions (or "bumps") increase surface area internal to the airfoil and provide
additional heat transfer effectiveness by disturbing airflow and "tripping" boundary
7
layer flow, increasing the exchange of energy (heat transfer), relative to a flat, nonenhanced surface. Heat transfer protrusions are applied only to a portion of airfoil
body, namely the area including and surrounding the leading edge to prevent
overheating downstream from the leading edge for narrower airfoils.
[0025] Referring to the drawings, FIG. 1 is a schematic view of an illustrative
turbomachine 90 in the form of a combustion turbine or gas turbine (GT) system
100 (hereinafter, “GT system 100”). GT system 100 includes a compressor 102
and a combustor 104. Combustor 104 includes a combustion region 105 and a head
end assembly 106 that includes one or more fuel nozzles. GT system 100 also
includes a turbine 108 and a common compressor/turbine shaft 110 (hereinafter
referred to as “rotor 110”). In one embodiment, GT system 100 is a 6F.03 FL18
engine, commercially available from General Electric Company, Greenville, S.C.
The present disclosure is not limited to any one particular GT system and may be
implanted in connection with other engines including, for example, the other HA,
F, B, LM, GT, TM and E-class engine models of General Electric Company and
engine models of other companies. Further, the teachings of the disclosure are not
necessarily applicable to only a GT system and may be applied to blades and/or
nozzles of other types of turbomachines, e.g., steam turbines, jet engines,
compressors, etc.
[0026] In operation, air flows through compressor 102, and compressed air is
supplied to combustor 104. Specifically, the compressed air is supplied to fuel
nozzles in head end assembly 106 that is integral to combustor 104. Head end
assembly 106 is in flow communication with combustion region 105. Fuel nozzles
in head end assembly 106 are also in flow communication with a fuel source (not
shown in FIG. 1), and the fuel nozzles channel fuel and air to combustion region
105. Combustor 104 ignites and combusts fuel to generate combustion products.
In the illustrative embodiment, there are a plurality of combustors 104 having head
end assemblies 106 with one or more fuel nozzles per head end assembly 106.
Combustors 104 are in flow communication with turbine 108 within which gas
stream thermal energy from the combustion products is converted to mechanical
rotational energy.
8
[0027] Turbine 108 is rotatably coupled to and drives rotor 110. Compressor 102
also is rotatably coupled to rotor 110. At least one end of rotor 110 may extend
axially away from turbine 108 and may be attached to a load or machinery (not
shown), such as, but not limited to, a generator, a load compressor, and/or another
turbine.
[0028] FIG. 2 shows a cross-section view of an illustrative portion of turbine 108
with four stages L0-L3 that may be used with GT system 100 in FIG. 1. The four
stages are referred to as L0, L1, L2, and L3. Stage L0 is the first stage and is the
smallest (in a radial direction) of the four stages. Stage L1 is the second stage after
the first stage in an axial direction. Stage L2 is the third stage and is the next stage
after the second stage in an axial direction. Stage L3 is the fourth, last stage in the
axial direction, and its blades are the largest (in a radial direction). It is to be
understood that four stages are shown as one example only, and each turbine may
have more or less than four stages.
[0029] A set of stationary vanes or nozzles 112 cooperate with a set of rotating
blades 114 to form each stage L0-L3 of turbine 108 and to define a portion of a flow
path through turbine 108. Rotating blades 114 in each set are coupled to a
respective rotor wheel 116 that couples them circumferentially to rotor 110 (FIG.
1). That is, a plurality of rotating blades 114 are mechanically coupled in a
circumferentially spaced manner to each rotor wheel 116. A static nozzle section
115 includes a plurality of stationary nozzles 112 circumferentially spaced around
rotor 110. Each nozzle 112 may include at least one endwall (or platform) 120, 122
connected with airfoil 130. In the example shown, nozzle 112 includes a radially
outer endwall 120 and a radially inner endwall 122. Radially outer endwall 120
couples nozzle(s) 112 to a casing 124 of turbine 108. In certain embodiments, static
nozzle section 115 is a second stage nozzle section, i.e., stage L1 in FIG. 2.
[0030] Turning to FIG. 3, a schematic perspective view of a cast turbine nozzle (or
simply, nozzle) 112 is shown, according to various embodiments to better illustrate
the parts of a nozzle. In FIG. 3, two nozzles 112 are shown as part of a static nozzle
section 115. In this manner, each nozzle 112 is a stationary nozzle, which forms
part of static nozzle section 115 (FIG. 2) and which forms part of an annulus of
9
stationary nozzles in a stage of a turbine (e.g., turbine 108), as previously described.
During operation of a turbine (e.g., turbine 108), nozzle 112 will remain stationary
in order to direct the flow of working fluid (e.g., gas, but could be steam) to one or
more movable blades (e.g., blades 114), causing those movable blades to initiate
rotation of a rotor 110. It is understood that nozzle 112 may be configured to couple
(mechanically couple via fasteners, welds, slot/grooves, etc.) with a plurality of
similar or distinct nozzles (e.g., nozzles 112 or other nozzles) to form an annulus of
nozzles in a stage L0-L3 of turbine 108.
[0001] Each turbine nozzle 112 can include a body 128 having an airfoil 130 having
a convex suction side 132, and a concave pressure side 134 (obstructed in FIG. 3)
opposing suction side 132. Nozzle 112 can also include a leading edge 136
spanning between pressure side 134 and suction side 132 and a trailing edge 138
opposing leading edge 136 and spanning between pressure side 134 and suction
side 132. As shown, and as previously noted, nozzle 112 can also include at least
one endwall 120, 122 (two shown) connected with airfoil(s) 130 along suction side
132, pressure side 134, trailing edge 138 and leading edge 136. In the example
shown, each nozzle 112 includes a radially outer endwall 120 and a radially inner
endwall 122. Radially outer endwalls 120 are configured to align on the radially
outer side of static nozzle section 115 (FIG. 2) and to couple respective nozzle(s)
112 to casing 124 (FIG. 2) of turbine 108 (FIG. 2). Radially inner endwalls 122 are
configured to align on the radially inner side of static nozzle section 115 (FIG. 2).
[0031] In various embodiments, each nozzle 112 includes a fillet 140, 142
connecting airfoil 130 and each respective endwall 120, 122. Fillet 140 can include
a weld or braze fillet, which may be formed via conventional metal-inert gas (MIG)
welding, tungsten-inert gas (TIG) welding, brazing, etc. Fillets 140, 142 can
overlap a portion of airfoil 130. The extent of overlap can vary from nozzle to
nozzle, stage to stage, and/or turbine to turbine.
[0032] Each nozzle 112 according to embodiments of the disclosure are cast, e.g.,
formed by molten material poured into a cast and hardened. Nozzle(s) 112 may
include any now known or later developed metal or metal alloy, such as a
superalloy, capable of withstanding the environment within turbine 108.
10
[0033] Each nozzle 112 may also include a cooling cavity 150 having an inner
surface 152 defined within body 128. FIG. 4 shows a perspective view of an
illustrative impingement insert or sleeve 154 that is inserted in each cooling cavity
150. That is, in operation, impingement sleeve 154 is positioned within cooling
cavity 150. As illustrated, impingement sleeve 154 includes a plurality of holes 156
therein configured to direct a coolant against inner surface 152 and about plurality
of heat transfer protrusions 160 (shown in FIGS. 5-7). As understood in the art,
cooling cavity 150 is fluidly coupled to a source of coolant, such as pressurized air
from compressor 102. The coolant passes through holes 156 in impingement insert
154 to impact against inner surface 152 to cool nozzle 112. Positioners 158 may
space impingement sleeve 154 from inner surface 152 to create an impingement
cooling area therebetween.
[0034] In certain commercial embodiments of turbine 108, it has been found
advantageous to scale nozzle 112 for use on turbines 108 of a different (e.g.,
smaller) gas turbine 100. Accordingly, the size of nozzle 112 (and particularly,
airfoil 130) is made smaller and/or narrower, which results in a radius of leading
edge 136 becoming increasingly smaller. The narrower airfoil 130 makes it more
difficult to cool leading edge 136 with conventional impingement cooling. For
example, turbine nozzle 112 may include a second stage nozzle for a 6-series gas
turbine.
[0035] Embodiments of the disclosure provide a plurality of heat transfer
protrusions 160 extending inwardly from inner surface 152 within body 128 in a
radially staggered columnar pattern. Protrusions 160 are integral with airfoil 130.
FIG. 5 shows a perspective view and FIG. 6 shows a slightly enlarged perspective
view of cast turbine nozzle(s) 112 including heat transfer protrusions 160, and FIG.
7 shows an enlarged perspective view of a plurality of heat transfer protrusions 160.
Heat transfer protrusions 160 extend from inner surface 152 at leading edge 136
along suction side 132 and along pressure side 134 in a radially staggered columnar
pattern. Heat transfer protrusions 160 do not extend along an entire chordal length
of each side 132, 134, as is conventional, because it has been discovered doing so
with narrower airfoils 130 causes overheating in the downstream areas, closer to
11
trailing edge 138. Rather, plurality of heat transfer protrusions extends in a range
of 28% to 32% of a camber length along suction side 132, and a range of 9% to
13% of the camber length along pressure side 134. “Camber length” represents a
distance from leading edge 136 to trailing edge 138 through a center of airfoil 130,
equidistant between suction side 132 and pressure side 134. A rough approximation
of a camber length CL is shown in FIG. 5. The extent of heat transfer protrusions
160 based on the stated percentages of camber length would be defined on each side
132, 134 at a location perpendicular to the camber length. In any event, only
portions of inner surface along each side 132, 134 are covered by heat transfer
protrusions 160, and inner surface 152 downstream of heat transfer protrusions 160
is devoid of protrusions or other structures that cause turbulence in the coolant flow
in the aft direction towards trailing edge 138. Heat transfer protrusions 160 may
extend to any radial extent on each side 132, 134 to achieve the desired heat transfer.
For example, they may span an entire radial length between endwalls 120, 122. In
contrast, in certain embodiments, heat transfer protrusions 160 may extend radially,
but stop in a range of 8 to 13 millimeters from one or more endwalls 120, 122.
[0036] FIG. 8 shows a plan view of inner surface 152 looking at tops of heat transfer
protrusions 160, and FIG. 9 shows a cross-sectional side view of heat transfer
protrusions 160 along line 9-9 in FIG. 8. As shown in FIGS. 8 and 9, inner surface
152 includes a planar surface 164 extending between adjacent heat transfer
protrusions 160. That is, planar surface 164 separates adjacent heat transfer
protrusions 160, with no inward or outward curvature of inner surface 152 other
than exists to form airfoil 130. In addition, as shown in FIG. 9, each heat transfer
protrusion 160 may have a frustoconical cross-section through a height thereof.
Each heat transfer protrusion 160 has an innermost surface 170 that is parallel with
inner surface 152 of cooling cavity 150 (FIGS. 5-6) between adjacent heat transfer
protrusions 160. As used herein, “innermost” indicates a portion of a structure
closest to a center of airfoil 130, and “outermost” indicates a portion of a structure
farthest from a center of airfoil 130. A height H of each heat transfer protrusion
160 from inner surface 152 of cooling cavity 150 to innermost surface 170 of heat
transfer protrusion 160 may range from 0.5 millimeters to 1.0 millimeters.
12
[0002] Heat transfer projections 160 may have an innermost width W1 that ranges
from 0.2 millimeters to 0.8 millimeters. Heat transfer projections 160 may have an
outermost width W2 that ranges from 0.6 millimeters to 1.2 millimeters. Outermost
width W2 is wider than innermost width W1. A ratio of innermost width W1 of
each heat transfer protrusion 160 to outermost width W2 of each heat transfer
protrusion 160 relative to inner surface 152 is in a range from 0.2 to 0.9. As shown
in FIG. 9, each heat transfer protrusion 160 may have a circular cross-section
through a width thereof. However, other non-elongated shapes may be possible.
Heat transfer projections 160 extend from inner surface 152 in a substantially
perpendicular angle α, i.e., at substantially 90°.
[0037] As shown in FIGS. 6 and 8, heat transfer protrusions 160 are arranged in a
radially staggered columnar pattern. As shown best in FIG. 8, the radially staggered
columnar pattern of the plurality of heat transfer protrusions 160 includes a plurality
of radially extending rows 176 (three shown in FIG. 8) that are radially staggered
(vertical on page) relative to one another. Any number of rows necessary to cover
the desired percentage of chordal length on each side 132, 134 may be used. A first
radial distance R1 between centers of heat transfer protrusions 160 in a same
radially extending row 176 may range from 0.9 millimeters to 1.4 millimeters. A
second radial distance R2 between centers of axially adjacent heat transfer
protrusions 160 in adjacent radially extending rows may range from 0.3 millimeters
to 0.9 millimeters. An axial distance AD between adjacent radially extending rows
176 of heat transfer protrusions 160 may range from 0.8 millimeters and 1.3
millimeters. An angular offset distance OF between heat transfer protrusions 160
may range, for example, from 0.9 millimeters to 1.4 millimeters. While a particular
radially staggered columnar pattern has been described herein, heat transfer
protrusions 160 may be arranged in alternative staggered columnar patterns to
achieve the desired heat transfer. In other embodiments, portions of innermost
widths W2 of adjacent heat transfer protrusions 160 may intersect or overlap.
[0038] In operation, coolant exits from impingement sleeve 154 (FIG. 4) and
impacts inner surface 152 of airfoil 130. Where present near leading edge 136, heat
transfer protrusions 160 cause turbulence in the coolant flow, increasing its heat
13
transfer capabilities. Heat transfer protrusions 160 may extend to any radial extent
and any chordal percentage to provide the desired heat transfer and cooling along
leading edge 136 and in areas of pressure side 134 and suction side 132 proximate
to leading edge 136.
[0039] Embodiments of the disclosure provide a cast turbine nozzle, a turbine
nozzle section and a turbine. The teachings are especially applicable to certain
second stage nozzles having smaller radius leading edges. The heat transfer
protrusions provide improved cooling effectiveness to maintain part life, turbine
efficiency, and power output for product specifications. More particularly, the heat
transfer protrusions or "bumps" increase surface area internal to the airfoil and
provide additional heat transfer effectiveness by disturbing airflow, increasing the
exchange of energy (heat transfer), relative to a flat, non-enhanced surface.
Because heat transfer protrusions are applied only to a portion of airfoil body, the
arrangement prevents overheating downstream from the leading edge for narrower
airfoils.
[0040] Approximating language, as used herein throughout the specification and
claims, may be applied to modify any quantitative representation that could
permissibly vary without resulting in a change in the basic function to which it is
related. Accordingly, a value modified by a term or terms, such as “about,”
“approximately” and “substantially,” are not to be limited to the precise value
specified. In at least some instances, the approximating language may correspond
to the precision of an instrument for measuring the value. Here and throughout the
specification and claims, range limitations may be combined and/or interchanged;
such ranges are identified and include all the sub-ranges contained therein unless
context or language indicates otherwise. “Approximately” as applied to a particular
value of a range applies to both end values, and unless otherwise dependent on the
precision of the instrument measuring the value, may indicate +/- 10% of the stated
value(s).
[0041] The corresponding structures, materials, acts, and equivalents of all means
or step plus function elements in the claims below are intended to include any
structure, material, or act for performing the function in combination with other
14
claimed elements as specifically claimed. The description of the present disclosure
has been presented for purposes of illustration and description but is not intended
to be exhaustive or limited to the disclosure in the form disclosed. Many
modifications and variations will be apparent to those of ordinary skill in the art
without departing from the scope and spirit of the disclosure. The embodiment was
chosen and described in order to best explain the principles of the disclosure and
the practical application, and to enable others of ordinary skill in the art to
understand the disclosure for various embodiments with various modifications as
are suited to the particular use contemplated.

WE CLAIM:

1. A cast turbine (108) nozzle (112), comprising:
an airfoil (130) having a body (128) including a suction side (132), a
pressure side (134) opposing the suction side (132), a leading edge (136) spanning
between the pressure side (134) and the suction side (132), a trailing edge (138)
opposing the leading edge (136) and spanning between the pressure side (134) and
the suction side (132), and a cooling cavity (150) defined by an inner surface (152)
of the body (128);
at least one endwall (120, 122) connected with the airfoil (130) along the
suction side (132), the pressure side (134), the trailing edge (138) and the leading
edge (136); and
a plurality of heat transfer protrusions (160) extending inwardly from the
inner surface (152) of the body (128) within the cooling cavity (150), the plurality
of heat transfer protrusions (160) extending from the leading edge (136) along the
suction side (132) and along the pressure side (134) in a radially staggered columnar
pattern,
wherein the inner surface (152) includes a planar surface (164) extending
between adjacent heat transfer protrusions (160).
2. The cast turbine (108) nozzle (112) of claim 1, wherein the turbine (108)
nozzle (112) includes a second stage nozzle (112).
3. The cast turbine (108) nozzle (112) of claim 1, wherein each heat transfer
protrusion (160) of the plurality of heat transfer protrusions (160) has a
frustoconical cross-section through a height thereof.
4. The cast turbine (108) nozzle (112) of claim 3, wherein each heat transfer
protrusion (160) of the plurality of heat transfer protrusions (160) has an innermost
surface (170) that is parallel with the inner surface (152) of the cooling cavity (150)
between adjacent heat transfer protrusions (160).
16
5. The cast turbine (108) nozzle (112) of claim 3, wherein each heat transfer
protrusion (160) has a circular cross-section through a width thereof.
6. The cast turbine (108) nozzle (112) of claim 1, further comprising an
impingement sleeve (154) within the cooling cavity (150), the impingement sleeve
(154) including a plurality of holes (156) therein configured to direct a coolant
against the inner surface (152) and about the plurality of heat transfer protrusions
(160).
7. The cast turbine (108) nozzle (112) of claim 1, wherein the at least one
endwall (120, 122) includes an inner endwall (122) or an outer endwall (120).
8. The cast turbine (108) nozzle (112) of claim 1, wherein a ratio of an
innermost width of each heat transfer protrusion (160) to an outermost width of
each heat transfer protrusion (160) relative to the inner surface (152) is in a range
from 0.2 to 0.9.
9. The cast turbine (108) nozzle (112) of claim 8, wherein the innermost width
of each heat transfer protrusion (160) ranges from 0.2 millimeters to 0.8
millimeters.
10. The cast turbine (108) nozzle (112) of claim 1, wherein a height of each heat
transfer protrusion (160) from the inner surface (152) ranges from 0.5 millimeters
to 1.0 millimeters.
11. The cast turbine (108) nozzle (112) of claim 1, wherein the radially
staggered columnar pattern of the plurality of heat transfer protrusions (160)
includes a plurality of radially extending rows (176) that are radially staggered
relative to one another, wherein a first radial distance between centers of heat
transfer protrusions (160) in a same radially extending row (176) ranges from 0.9
17
millimeters to 1.4 millimeters, and a second radial distance between centers of
axially adjacent heat transfer protrusions (160) in adjacent radially extending rows
(176) ranges from 0.3 millimeters to 0.9 millimeters, and
wherein an axial distance between adjacent radially extending rows (176)
of the heat transfer protrusions (160) ranges from 0.8 millimeters and 1.3
millimeters.
12. The cast turbine (108) nozzle (112) of claim 1, wherein the plurality of heat
transfer protrusions (160) extends in a range of 28% to 32% of a camber length
along suction side (132), and a range of 9% to 13% of the camber length along
pressure side (134).
13. A nozzle (112) section for a turbine (108), the nozzle (112) section
comprising:
a set of nozzles (112), the set of nozzles (112) including at least one cast
nozzle (112) having:
an airfoil (130) having a body (128) including a suction side (132), a
pressure side (134) opposing the suction side (132), a leading edge (136) spanning
between the pressure side (134) and the suction side (132), a trailing edge (138)
opposing the leading edge (136) and spanning between the pressure side (134) and
the suction side (132), and a cooling cavity (150) having an inner surface (152)
defined within the body (128);
at least one endwall (120, 122) connected with the airfoil (130) along the
suction side (132), the pressure side (134), the trailing edge (138) and the leading
edge (136); and
a plurality of heat transfer protrusions (160) extending inwardly from the
inner surface (152) within the body (128), the plurality of heat transfer protrusions
(160) extending from the leading edge (136) along the suction side (132) and along
the pressure side (134) in a radially staggered columnar pattern,
wherein the inner surface (152) includes a planar surface (164) extending
between adjacent heat transfer protrusions (160).
18
14. The nozzle (112) section of claim 13, wherein the static nozzle (112) section
is a second stage nozzle (112) section.
15. The nozzle (112) section of claim 13, wherein each heat transfer protrusion
(160) of the plurality of heat transfer protrusions (160) has a frustoconical crosssection through a height thereof.