U.S. patent application number 15/811862 was filed with the patent office on 2019-01-24 for methods for manufacturing a turbine nozzle with single crystal alloy nozzle segments.
This patent application is currently assigned to HONEYWELL INTERNATIONAL INC.. The applicant listed for this patent is HONEYWELL INTERNATIONAL INC.. Invention is credited to Ardeshir Riahi, Jason Smoke, Bradley Reed Tucker, Ed Zurmehly.
Application Number | 20190022781 15/811862 |
Document ID | / |
Family ID | 54007466 |
Filed Date | 2019-01-24 |
United States Patent
Application |
20190022781 |
Kind Code |
A1 |
Tucker; Bradley Reed ; et
al. |
January 24, 2019 |
METHODS FOR MANUFACTURING A TURBINE NOZZLE WITH SINGLE CRYSTAL
ALLOY NOZZLE SEGMENTS
Abstract
Methods for manufacturing a turbine nozzle are provided. A
plurality of nozzle segments is formed. Each nozzle segment
comprises an endwall ring portion with at least one vane. The
plurality of nozzle segments are connected to an annular endwall
forming a segmented annular endwall concentric to the annular
endwall with the at least one vane of each nozzle segment extending
between the segmented annular endwall and the annular endwall.
Inventors: |
Tucker; Bradley Reed;
(Chandler, AZ) ; Riahi; Ardeshir; (Scottsdale,
AZ) ; Smoke; Jason; (Phoenix, AZ) ; Zurmehly;
Ed; (Phoenix, AZ) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
HONEYWELL INTERNATIONAL INC. |
Morris Plains |
NJ |
US |
|
|
Assignee: |
HONEYWELL INTERNATIONAL
INC.
Morris Plains
NJ
|
Family ID: |
54007466 |
Appl. No.: |
15/811862 |
Filed: |
November 14, 2017 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
14341162 |
Jul 25, 2014 |
9844826 |
|
|
15811862 |
|
|
|
|
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F05D 2230/21 20130101;
Y10T 29/49346 20150115; B23K 1/0018 20130101; F05D 2230/90
20130101; F05D 2230/237 20130101; B23P 15/006 20130101; F01D 9/044
20130101; B22D 19/0081 20130101; F05D 2300/607 20130101; B22D 19/04
20130101 |
International
Class: |
B23K 1/00 20060101
B23K001/00; F01D 9/04 20060101 F01D009/04; B23P 15/00 20060101
B23P015/00; B22D 19/04 20060101 B22D019/04; B22D 19/00 20060101
B22D019/00 |
Claims
1. A method for manufacturing a turbine nozzle comprising: forming
a plurality of nozzle segments, each nozzle segment comprising an
endwall ring portion with at least one vane, the at least one vane
having a free end portion; connecting the free end portion of the
at least one vane of each nozzle segment of the plurality of nozzle
segments to an annular endwall, with the endwall ring portion of
each nozzle segment of the plurality of nozzle segments forming a
segmented annular endwall concentric to the annular endwall with
the at least one vane of each nozzle segment extending between the
segmented annular endwall and the annular endwall; and forming the
annular endwall prior to the connecting by separately casting the
annular endwall as one-piece, wherein the step of connecting the
plurality of nozzle segments comprises brazing the free end portion
of the at least one vane of each nozzle segment to the annular
endwall.
2. The method of claim 1, wherein the step of forming a plurality
of nozzle segments comprises forming the plurality of nozzle
segments with a single crystal material.
3. The method of claim 1, wherein the step of forming a plurality
of nozzle segments comprises forming by casting.
4. The method of claim 1, wherein the step of forming a plurality
of nozzle segments comprises forming a plurality of singlet nozzle
segments, doublet nozzle segments, triplet nozzle segments,
quadruplet nozzle segments, or combinations thereof.
5. The method of claim 1, further comprising the step of processing
at least one nozzle segment of the plurality of nozzle segments
after the forming the plurality of nozzle segments and prior to the
connecting step, wherein the step of processing comprises applying
a protective coating to at least one nozzle segment of the
plurality of nozzle segments, the annular endwall, or both.
6. The method of claim 1, further comprising the step of processing
at least one nozzle segment of the plurality of nozzle segments
prior to the connecting step, wherein the step of processing
comprises machining at least one cooling hole in at least one
nozzle segment, machining a feather-seal slot in opposing ends of
at least two nozzle segments for receiving an interlocking feather
seal, or both.
7. The method of claim 1, wherein the step of connecting the
plurality of nozzle segments comprises circumferentially spacing
the plurality of nozzle segments along the annular endwall with the
endwall ring portion of each nozzle segment cooperating to form the
segmented annular endwall.
8. (canceled)
9. The method of claim 6, further comprising the step of sealing
the plurality of nozzle segments together during or after the
connecting step, the step of sealing comprising inserting the
interlocking feather seal into intersecting feather-seal slots
machined into opposed end faces of the endwall ring portion of at
least a pair of circumferentially adjacent nozzle segments.
10. A method for manufacturing a turbine nozzle comprising: forming
a plurality of nozzle segments configured to be connected to an
outer annular endwall, each nozzle segment integrally cast as one
piece from a single crystal alloy material and comprised of an
endwall ring portion and at least one vane, the at least one vane
having a free end portion; brazing the free end portion of the at
least one vane of each nozzle segment to the outer annular endwall,
the endwall ring portion of each nozzle segment cooperating with
the endwall ring portion of a circumferentially adjacent nozzle
segment to form a segmented inner annular endwall concentric to the
outer annular endwall with the at least one vane of each nozzle
segment extending between the segmented inner annular endwall and
the outer annular endwall, the brazing of the free end portion of
the at least one vane of each nozzle segment to the outer annular
endwall forming the turbine nozzle; and forming the outer annular
endwall prior to the brazing by separately casting the outer
annular endwall as a one-piece ring.
11. The method of claim 10, wherein the step of forming a plurality
of nozzle segments comprises forming a plurality of singlet nozzle
segments, doublet nozzle segments, triplet nozzle segments,
quadruplet nozzle segments, or combinations thereof.
12. The method of claim 10, further comprising the step of
processing at least one nozzle segment of the plurality of nozzle
segments after forming the plurality of nozzle segments and prior
to brazing, wherein the step of processing comprises applying a
protective coating to the at least one nozzle segment.
13. The method of claim 10, further comprising the step of
processing at least one nozzle segment of the plurality of nozzle
segments prior to brazing, wherein the step of processing comprises
machining at least one cooling hole in the at least one nozzle
segment.
14. The method of claim 10, further comprising the step of
processing at least one nozzle segment of the plurality of nozzle
segments prior to brazing, wherein the step of processing comprises
machining a feather-seal slot in opposing end faces of the endwall
ring portion of at least two nozzle segments for receiving an
interlocking feather seal.
15. The method of claim 10, further comprising the step of sealing
the plurality of nozzle segments together during or after the
connecting step, the step of sealing comprising inserting the
interlocking feather seal into intersecting feather-seal slots
machined into at least a pair of circumferentially adjacent nozzle
segments.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application is a divisional of U.S. patent application
Ser. No. 14/341,162 filed on Jul. 25, 2014. The relevant disclosure
of the above application is incorporated herein by reference.
TECHNICAL FIELD
[0002] The present invention generally relates to gas turbine
engines, and more particularly relates to methods for manufacturing
a turbine nozzle with single crystal alloy nozzle segments.
BACKGROUND
[0003] Gas turbine engines may be used to power various types of
vehicles and systems, such as air or land-based vehicles. In
typical gas turbine engines, compressed air generated by axial
and/or radial compressors is mixed with fuel and burned, and the
expanding hot combustion gases are directed along a flowpath and
through a turbine nozzle having stationary vanes. The combustion
gas flow deflects off of the vanes and impinges upon turbine blades
of a turbine rotor. A rotatable turbine disk or wheel, from which
the turbine blades extend, spins at high speeds to produce power.
Gas turbine engines used in aircraft use the power to draw more air
into the engine and to pass high velocity combustion gas out of the
gas turbine aft end to produce a forward thrust. Other gas turbine
engines may use the power to turn a propeller or an electrical
generator.
[0004] Typically, the stationary vanes of the turbine nozzle extend
between an inner endwall ring (also known as a "hub ring") and an
outer endwall ring (also known as a "shroud ring"). The inner and
outer endwall rings define a portion of the flowpath along which
the combustion gas travels. In some cases, the inner and/or outer
endwalls rings are initially formed as segments, and the segments
are subsequently assembled together to form a full ring (a
"conventional segmented turbine nozzle"). Conventional segmented
turbine nozzles may experience significant leakage where the
adjacent segments meet at segment platform seal gaps and
intermittent flange surfaces. Additionally, high leakage may exist
where the segments mate to the supporting structure due to
dimensional variation caused by individually machined segments. The
leakage between segments is detrimental to the gas turbine engine
in two major ways. First, the leakage increases chargeable cooling
flow that does not get turned by the turbine nozzle to produce work
across the turbine rotor, thus increasing fuel consumption.
Secondly, the increased leakage flow wastes precious cooling flow
that could be used for combustor and turbine component cooling. As
combustor and turbine nozzle distress are among the top
contributors to hot section replacement overhaul costs, gas turbine
engine designers are eagerly seeking ways to reduce this
detrimental leakage in segmented turbine nozzles and use the flow
to cool the combustor and nozzle instead, thereby improving
component durability and service life.
[0005] Turbine nozzles may also be manufactured by bi-casting the
stationary turbine vanes with the inner and outer endwall rings, so
that the rings and the vanes comprise a single, unitary turbine
nozzle (a "conventional bi-cast turbine nozzle"). Though bi-cast
inner and outer endwall rings reduce turbine nozzle leakage, they
may be difficult and/or time consuming to manufacture, with reduced
manufacturing yields. For example, a bi-cast turbine nozzle may
suffer cracking distress and reduced service life due to
thermo-mechanical fatigue (TMF) caused by a lack of radial
compliance between the vanes and endwall rings. In addition,
bi-casting of the endwall rings requires that the endwall rings be
fabricated from an equi-axed alloy that has lower strength and
oxidation capabilities than a single crystal alloy. Moreover,
protective coatings may be relatively difficult to apply to
conventional bi-cast turbine nozzles. The coated surfaces of
conventional bi-cast turbine nozzles show irregularities on the
surfaces where "shadows" cast by adjacent vanes result in a
non-optimal coating microstructure and thickness distribution.
[0006] Hence, there is a need for improved methods for
manufacturing a turbine nozzle with single crystal alloy nozzle
segments. There is also a need for improved methods for
manufacturing a turbine nozzle with single crystal alloy nozzle
segments to reduce leakage and improve coating application.
Furthermore, other desirable features and characteristics of the
inventive subject matter will become apparent from the subsequent
detailed description of the inventive subject matter and the
appended claims, taken in conjunction with the accompanying
drawings and this background of the inventive subject matter.
BRIEF SUMMARY
[0007] Methods are provided for manufacturing a turbine nozzle. In
accordance with one exemplary embodiment, the method comprises
forming a plurality of nozzle segments. Each nozzle segment
comprises an endwall ring portion with at least one vane extending
therefrom. The plurality of nozzle segments are connected to an
annular endwall forming a segmented annular endwall concentric to
the annular endwall with the at least one vane of each nozzle
segment extending between the segmented annular endwall and the
annular endwall.
[0008] Methods are provided for manufacturing a turbine nozzle, in
accordance with yet another exemplary embodiment of the present
invention. The method comprises forming a plurality of nozzle
segments configured to be connected to an annular endwall. Each
nozzle segment is integrally cast as one piece from a single
crystal alloy material and is comprised of an endwall ring portion
and at least one vane having a free end portion. The annular
endwall is bi-cast around the plurality of nozzle segments. The
endwall ring portion of each nozzle segment cooperates with a
circumferentially adjacent nozzle segment to form a segmented
annular endwall. The at least one vane of each nozzle segment
extends between the segmented annular endwall and the annular
endwall.
[0009] Methods are provided for manufacturing a turbine nozzle, in
accordance with yet another exemplary embodiment of the present
invention. The method comprises forming a plurality of nozzle
segments configured to be connected to an annular endwall. Each
nozzle segment is integrally cast as one piece from a single
crystal alloy material and is comprised of an endwall ring portion
and at least one vane. The at least one vane of each nozzle segment
is brazed to the annular endwall. The endwall ring portion of each
nozzle segment cooperates with the endwall ring portion of a
circumferentially adjacent nozzle segment to form a segmented
annular endwall. The segmented annular endwall is concentric to the
annular endwall with the at least one vane of each nozzle segment
extending between the segmented annular endwall and the annular
endwall.
[0010] Furthermore, other desirable features and characteristics of
the methods will become apparent from the subsequent detailed
description and the appended claims, taken in conjunction with the
accompanying drawings and the preceding background.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] The present invention will hereinafter be described in
conjunction with the following drawing figures, wherein like
numerals denote like elements, and wherein:
[0012] FIG. 1 is a cross-sectional side view of a turbine section
of an exemplary gas turbine engine;
[0013] FIG. 2 is a simplified side view of a portion of a turbine
nozzle manufactured according to exemplary embodiments, the turbine
nozzle comprising an outer endwall ring disposed concentrically to
and radially outwardly from an inner endwall ring and a plurality
of stationary vanes extending between the inner and outer endwall
rings;
[0014] FIG. 3 is a flow diagram of a method for manufacturing a
turbine nozzle with single crystal nozzle segments, according to
exemplary embodiments of the present invention;
[0015] FIG. 4 is an isometric view of a cast single crystal nozzle
segment (a singlet) of the turbine nozzle of FIG. 2, according to
exemplary embodiments;
[0016] FIG. 5 is an isometric view of the cast single crystal
nozzle segment of FIG. 4 including a feather-seal slot machined
into opposing end faces thereof (only one of which is shown);
[0017] FIG. 6 is a cross-sectional side view of a portion of the
turbine nozzle of FIG. 2 taken along line A-A thereof, depicting
singlet nozzle segments bi-cast with the outer endwall ring as
indicated by dotted circle B and the circumferentially adjacent
nozzle segments sealed together by a feather seal received in
intersecting feather-seal slots according to exemplary embodiments
of the present invention;
[0018] FIG. 7 is a cross-sectional side view similar to FIG. 6 of a
portion of the turbine nozzle of FIG. 2, depicting a doublet nozzle
segment circumferentially adjacent and between a pair of nozzle
segments (partially shown); and
[0019] FIG. 8 is a cross-sectional view of a portion of a turbine
nozzle similar to FIG. 2 taken along line A-A thereof, manufactured
according to an alternative exemplary embodiment depicting a free
end portion of a stationary vane of a nozzle segment (partially
shown) brazed into the outer endwall ring of the turbine nozzle at
braze joints A.
DETAILED DESCRIPTION
[0020] The following detailed description is merely exemplary in
nature and is not intended to limit the invention or the
application and uses of the invention. As used herein, the word
"exemplary" means "serving as an example, instance, or
illustration." Thus, any embodiment described herein as "exemplary"
is not necessarily to be construed as preferred or advantageous
over other embodiments. All of the embodiments described herein are
exemplary embodiments provided to enable persons skilled in the art
to make or use the invention and not to limit the scope of the
invention which is defined by the claims. Furthermore, there is no
intention to be bound by any expressed or implied theory presented
in the preceding technical field, background, brief summary, or the
following detailed description.
[0021] Various embodiments are directed to methods for
manufacturing a turbine nozzle with single crystal alloy nozzle
segments. Such manufacturing methods enable using a single crystal
alloy material, thereby conferring increased strength and/or
oxidation life to the turbine nozzle manufactured according to
exemplary embodiments as described herein. In addition, such
methods also improve manufacturing yields and protective coating
characteristics of the turbine nozzle, as the nozzle segments may
be coated prior to assembly as hereinafter described, providing an
improved protective coating microstructure and thickness
distribution. Such methods also enable the manufacture of a turbine
nozzle with radial compliance without the use of a slip joint and
reduced leakage flows relative to conventional segmented turbine
nozzles.
[0022] FIG. 1 is a cross-sectional side view of a portion of a
turbine section 100 of a gas turbine engine. The turbine section
100 receives high temperature (e.g., a temperature typically in the
range of about 1100 to 1800.degree. C.) gases from an upstream
engine combustor (not shown) to produce energy for the engine
and/or components coupled to the engine. The turbine section 100
includes a turbine nozzle 104 that has a plurality of static vanes
106 (only one of which is shown) that direct the gases from the
combustor to a turbine rotor 110. According to an embodiment, the
turbine rotor 110 includes a plurality of blades 112 (only one of
which is shown) that are retained in axial position by a retention
plate 116. When the blades 112 are impinged upon by the gases, the
gases cause the turbine rotor 110 to spin. According to an
embodiment, an outer circumferential wall 118 (also referred to
hereinafter as "an outer endwall ring" 206) surrounds the static
vanes 106 and defines a portion of a plenum 120. The plenum 120
receives air from a compressor section (not shown), which may be
directed through one or more openings in the outer circumferential
wall 118 to cool the static vanes 106.
[0023] FIG. 2 is a side view of a portion of the turbine nozzle 104
manufactured according to exemplary embodiments of the present
invention. The turbine nozzle 104 includes the plurality of static
or stationary vanes (only one of which is shown) 106, an inner
endwall ring 204, and an outer endwall ring 206. The vanes 106
extend between the rings 204, 206 and, as hereinafter described,
may form a joint with the outer endwall ring 206 (e.g., such as by
bi-casting as depicted in FIG. 2 or by braze joint A as depicted in
FIG. 8). The inner endwall ring 204 is disposed concentric to and
radially inwardly from the outer endwall ring 206. As hereinafter
described, one of the inner endwall ring or the outer endwall ring
comprises an "annular endwall" and the other of the inner endwall
ring or the outer endwall ring comprises a "segmented annular
endwall" in the turbine nozzle. In the depicted embodiments of
FIGS. 6 through 8, the inner endwall ring comprises the segmented
annular endwall and the outer endwall ring comprises the annular
endwall. In an alternative embodiment (not shown), the outer
endwall ring comprises the segmented annular endwall and the inner
endwall ring comprises the annular endwall, in which case the vanes
form a joint with the inner endwall ring.
[0024] Each vane 106 includes an airfoil 212, an inner end 208, and
an outer end 210. The airfoil has two outer walls 218 (only one of
which is shown), each having outer surfaces that define an airfoil
shape. The airfoil shape includes a leading edge 220, a trailing
edge 222, a pressure side 224 along the first outer wall 218, and a
suction side along the second outer wall (not shown). In some
embodiments, though not shown, the vane 106 may have an internal
cooling circuit formed therein, that may extend from an opening in
the first end through the vane and may include various passages
that eventually communicate with trailing edge openings or other
openings (not shown) that may be formed in the vane.
[0025] Referring now to FIG. 3, according to exemplary embodiments,
a method 10 for manufacturing a turbine nozzle 104 such as depicted
in FIGS. 1 and 2 begins by forming a plurality of nozzle segments
300 (step 20) (as exemplified by nozzle segments 300 depicted in
FIGS. 4 through 7). The plurality of nozzle segments may be formed,
for example, by casting. Each nozzle segment 300 comprises at least
one vane 106 supported by a corresponding endwall ring portion 302.
The endwall ring portions 302, when the nozzle segments 300 are
assembled as hereinafter described, make up the segmented annular
endwall. The endwall ring portions may also be referred to herein
as "vane platforms." The nozzle segment 300 may be a singlet
(single vane) 300a as best depicted in FIGS. 4, 5, and 6 or a
multiplet such as a doublet (two vanes) 300b as depicted in FIG. 7,
a triplet (three vanes) (not shown), or a quadruplet (four vanes)
(not shown), etc. In general, singlets and doublets are preferred
but nozzle segments with greater than two vanes may be used to form
the segmented annular endwall. The coating advantages as
hereinafter described decline with the number of vanes in each
nozzle segment, however the leakage flows also decrease. While FIG.
6 depicts assembled singlet nozzle segments 300a, it is to be
understood that the turbine nozzle may be manufactured with
combinations of singlets, doublets, triplets, quadruplets, etc. In
addition, while the nozzle segments depicted in FIGS. 4 through 8
include endwall ring portions of the inner endwall ring, it is to
be understood that the nozzle segments may alternatively comprise
endwall ring portions of the outer endwall ring and the at least
one vane.
[0026] The nozzle segments 300 may be cast of a single crystal
alloy. As used herein, a "single crystal alloy" is an alloy in
which substantially the entire alloy has a single crystallographic
orientation, without the presence of high angle grain boundaries. A
small amount of low angle grain boundaries such as tilt, or twist
boundaries are permitted within such a single crystal article but
are preferably not present. However, such low angle boundaries are
often present after solidification and formation of the single
crystal article, or after some deformation of the article during
creep or other light deformation process. Other minor
irregularities are also permitted within the scope of the term
"single crystal". For example, small areas of high angle grain
boundaries may be formed in various portions of the article, due to
the inability of the single crystal to grow perfectly near corners
and the like. Such deviations from a perfect single crystal, which
are found in normal commercial production operations, are within
the scope of the term single crystal as used herein. The nozzle
segments may be cast by methods well known in the art.
[0027] Referring again to FIG. 3, according to exemplary
embodiments, method 10 continues by optionally processing one or
more of the nozzle segments (step 30). In step 30, the nozzle
segments may undergo processing including applying at least one
protective coating, machining as hereinafter described, or
combinations thereof. Such processing may be advantageously
performed after forming the plurality of nozzle segments and prior
to connecting the nozzle segments together as hereinafter
described. Such processing may additionally or alternatively be
used at other times.
[0028] The at least one protective coating may be, for example, a
bond coating, a thermal barrier coating (TBC), an oxidation
resistant coating, or the like. The ability to apply at least one
protective coating to a nozzle segment, prior to connecting the
nozzle segments improves the coating microstructure and thickness
distribution because there are fewer or no adjacent vanes (no
adjacent vanes in the case of a singlet nozzle segment) to cause
surface shadowing. As noted previously, shadows cast by adjacent
vanes place adjacent surfaces in the shadows and threaten the
coating microstructure and thickness distribution.
[0029] The optional processing step 20 may alternatively or
additionally comprise machining one or more of the nozzle segments
to include features such as cooling holes (not shown), feather-seal
slots 304 (FIG. 5), etc. Opposing end faces 306 of the endwall ring
portions of the nozzle segments 300 may be machined to incorporate
the feather-seal slots 304, as depicted in FIGS. 5, 6, and 7. The
feather-seal slots 304 are configured to receive a feather seal 308
(FIGS. 6 and 7) in an optional sealing step (step 50) as
hereinafter described. If the nozzle segments are not to be sealed
together, no feather-seal slots are necessary. The forming (step
20) with or without optional processing (step 30) of the plurality
of nozzle segments results in "prefabricated nozzle segments" or
simply "nozzle segments."
[0030] Referring again to FIGS. 3 and 6 through 8, according to
exemplary embodiments, the method 10 for manufacturing a turbine
nozzle with single crystal nozzle segments continues by providing
the one-piece annular endwall (step 35). In an embodiment, the step
of providing the one-piece annular endwall comprises bi-casting the
one-piece annular endwall. Bi-casting is well known in the art. The
material for the annular endwall (the outer endwall ring 206 in
FIGS. 6 and 7) is cast in a casting mold around the free end
portion of the vanes of the prefabricated nozzle segments to both
provide the annular endwall (step 35) and connect the annular
endwall with the nozzle segments (step 40). Thus, for the bi-cast
embodiment, steps 35 and 40 are performed simultaneously during
bi-casting. The term "bi-casting" refers to the method in which the
prefabricated nozzle segments are disposed in the casting mold. A
"dog-bone feature" 82 (FIGS. 2 and 6 through 7) at the free end
portion of the at least one vane of each nozzle segment fixedly
attaches the vanes of the prefabricated nozzle segments to the
bi-cast annular endwall. While bi-casting itself connects the
nozzle segments with the annular endwall, the dog-bone feature 82
provides a mechanical bond with the prefabricated nozzle segment.
FIGS. 6 and 7 depict the plurality of nozzle segments both
connected and fixedly attached to the annular endwall formed by
bi-casting, thereby forming a turbine nozzle assembly.
[0031] In an alternative embodiment, the step of providing the
one-piece annular endwall (step 35) comprises separately casting
the one-piece annular endwall. If the annular endwall is separately
cast, the separately cast annular endwall is then connected to the
plurality of nozzle segments in the connecting step 40 (thereby
forming a turbine nozzle assembly) by brazing the free end portion
of the vanes of the plurality of prefabricated nozzle segments to
the annular endwall forming braze joints A (FIG. 8). The turbine
nozzle manufactured according to the alternative embodiment appears
similar to the turbine nozzle depicted in FIG. 2 except there is no
"dog-bone feature" 82 as in FIGS. 2 and 6 through 7 at the free end
portion of the vanes of the nozzle segments and access to the
annular endwall is provided for applying a braze material to form
the braze joints A.
[0032] Connecting the annular endwall to the plurality of nozzle
segments in the connecting step 40 forms the turbine nozzle
assembly comprising the segmented annular endwall concentric to the
annular endwall with the at least one vane of each nozzle segment
extending between the segmented annular endwall and the annular
endwall. The single crystal nozzle segments of the turbine nozzle
assembly are arranged in a circumferentially spaced relation. The
endwall ring portions of circumferentially adjacent nozzle segments
are butted end to end forming the segmented annular endwall. The
prefabricated nozzle segments in the segmented annular endwall are
discrete, without any direct connection to each other. As the
segmented annular endwall comprises the endwall ring portions of
the plurality of nozzle segments, substantially free radial
movement is ensured by the discrete nozzle segments without the
need of a slip joint at the interface of the vanes and one of the
endwalls as in conventional bi-cast turbine nozzles. Manufacturing
yields of turbine nozzles may also be increased because defective
nozzle segments can be scrapped if found to be defective, rather
than scrapping an entire turbine nozzle.
[0033] Suitable exemplary material for the annular endwall
comprises an equi-axed alloy or other materials that do not have a
single crystal orientation. For example, the one-piece annular
endwall ring may be formed of metal or other materials that can
withstand the extremely high operating temperatures (greater than
about 2800.degree. Fahrenheit) to which they are exposed in the gas
turbine engine. The one-piece annular ring may be a unitary cast
alloy structure produced by a precision casting operation utilizing
various superalloy compositions. Various types of alloy, superalloy
compositions and manufactures of such compositions are known to
those skilled in the art.
[0034] Referring again to FIG. 3, according to exemplary
embodiments, the method 10 for manufacturing the turbine nozzle
with single crystal nozzle segments continues by optionally sealing
the plurality of prefabricated nozzle segments together in the
turbine nozzle assembly (step 50). The optional sealing step
provides a direct connection between the prefabricated nozzle
segments of the turbine nozzle assembly. A feather seal 308 fits
into the intersecting feather-seal slots 304 in each endwall ring
portion (i.e., the vane platform) so as to seal the interface
between abutting nozzle segments that make up the segmented annular
endwall. The purpose of the feather seals is to seal off the fluid
working medium from the ambient surrounding the fluid working
medium and vice versa. For the bi-cast turbine nozzle assembly, the
feather seals may be assembled prior to bi-casting as previously
described, and thus trapped in-place by the slot geometry, or may
be assembled after bi-casting and retained by a separate feather
seal member.
[0035] To seal between the nozzle segments, the edge of the feather
seal may be inserted into the feather-seal slot 304 of one of the
prefabricated nozzle segments. The next adjacent nozzle segment is
aligned so that its complementary feather-seal slot 304 aligns with
the opposite edge of the feather seal. The nozzle segments are
urged toward each other so that they are in abutting end to end
position. The adjacent nozzle segments 300 interlock to form a
plurality of contiguous circumferentially adjacent nozzle segments
as depicted in FIGS. 6 and 7. The dimensions (e.g., the widths and
heights) and surface contours of the nozzle segments 300 are
preferably substantially identical at the interlocking interface to
provide a continuous or uninterrupted transition between nozzle
segment surfaces to minimize leakage. As a result of this
structural configuration, the turbine nozzle provides a
substantially uninterrupted flowpath to minimize leakage of
combustion gas. FIGS. 6 and 7 are cross-sectional side views
depicting a portion of the turbine nozzle 104/turbine nozzle
assembly (FIG. 2) after the sealing step 50. More specifically,
FIG. 6 depicts a singlet nozzle segment 300a interlocking with
other singlet nozzle segments 300a that are circumferentially
adjacent. FIG. 7 depicts a doublet nozzle segment 300b interlocking
with circumferentially adjacent nozzle segments 300 (partially
shown).
[0036] Referring again to FIG. 3, in accordance with exemplary
embodiments, method 10 for manufacturing a turbine nozzle with
single crystal nozzle segments continues by optionally finishing
the turbine nozzle assembly (step 60). Complete finish machining of
the turbine nozzle assembly substantially ensures precise machining
of critical sealing surfaces, thereby further reducing leakage
flows, particularly between the turbine nozzle and mating structure
that exists in conventional segmented turbine nozzles. The
exemplary turbine nozzle illustrated partially in FIG. 2 is a
turbine nozzle in which the critical sealing surfaces have been
finished by machining or the like. In some exemplary embodiments,
no such finishing treatments are necessary and step 60 may be
omitted.
[0037] From the foregoing, it is to be appreciated that the methods
for manufacturing turbine nozzles with single crystal alloy nozzle
segments are provided. Such methods enable manufacturing a turbine
nozzle with much lower technical risk relative to conventional
bi-cast turbine nozzles as radial compliance is ensured without
dependence on manufacture of a slip joint. Manufacture of turbine
nozzles with single crystal alloy nozzle segments also permits
turbine nozzles with less leakage flows and thus lower cooling
flows than conventional segmented turbine nozzles.
[0038] While at least one exemplary embodiment has been presented
in the foregoing detailed description of the invention, it should
be appreciated that a vast number of variations exist. It should
also be appreciated that the exemplary embodiment or exemplary
embodiments are only examples, and are not intended to limit the
scope, applicability, or configuration of the invention in any way.
Rather, the foregoing detailed description will provide those
skilled in the art with a convenient road map for implementing an
exemplary embodiment of the invention. It being understood that
various changes may be made in the function and arrangement of
elements described in an exemplary embodiment without departing
from the scope of the invention as set forth in the appended
claims.
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