U.S. patent application number 09/916352 was filed with the patent office on 2003-06-19 for selective step turbine nozzle.
Invention is credited to Ackerman, Robert Ingram, Heyward, John Peter, White, Gregory Alan.
Application Number | 20030113206 09/916352 |
Document ID | / |
Family ID | 25437131 |
Filed Date | 2003-06-19 |
United States Patent
Application |
20030113206 |
Kind Code |
A1 |
Heyward, John Peter ; et
al. |
June 19, 2003 |
SELECTIVE STEP TURBINE NOZZLE
Abstract
A turbine nozzle includes segments of outer and inner bands
supporting corresponding vane pairs. The bands adjoin each other at
corresponding ends along splitlines, with each band having a
forward land, an opposite aft land, and a middle land extending
therebetween. The forward lands have a nominal aft-facing step, the
aft lands have a nominal forward-facing step and the middle lands
are nominally flush.
Inventors: |
Heyward, John Peter;
(Loveland, OH) ; Ackerman, Robert Ingram; (West
Chester, OH) ; White, Gregory Alan; (Cincinnati,
OH) |
Correspondence
Address: |
FRANCIS L. CONTE, ESQ.
6 PURITAN AVENUE
SWAMPSCOTT
MA
01907
US
|
Family ID: |
25437131 |
Appl. No.: |
09/916352 |
Filed: |
July 27, 2001 |
Current U.S.
Class: |
415/189 |
Current CPC
Class: |
Y10T 29/49323 20150115;
F05D 2230/23 20130101; Y02T 50/60 20130101; F05D 2230/21 20130101;
F01D 9/041 20130101; F01D 5/143 20130101; Y02T 50/673 20130101 |
Class at
Publication: |
415/189 |
International
Class: |
F01D 001/02 |
Claims
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 in which we claim:
1. A turbine nozzle comprising: a row of nozzle segments each
including a pair of vanes mounted at opposite ends to outer and
inner bands; each of said vanes having axially opposite leading and
trailing edges and circumferentially opposite pressure and suction
sides extending radially between said opposite ends; each of said
bands having circumferentially opposite first and second ends
adjoining at respective splitlines therebetween; each of said band
ends having a forward land extending along said splitline near said
vane leading edge, an aft land extending along said splitline near
said vane trailing edge, and a middle land extending along said
splitline over an intermediate portion of said vane between said
forward and aft lands to collectively form flowpath surfaces
bounding combustion gas flow between said vanes; and said forward
lands at said splitlines having a nominal aft-facing, said aft
lands at said splitlines having a nominal forward-facing step and
said middle lands being nominally flush.
2. A nozzle according to claim 1 wherein said forward and aft lands
blend with said middle lands for blending said aft-facing steps and
forward-facing steps with said nominally flush middle lands.
3. A nozzle according to claim 2 wherein said middle lands are
sized in length along said splitlines to prevent streamlines of
combustion gases from flowing into said aft-facing steps at said
forward lands, and from flowing into said forward-facing steps at
said aft lands.
4. A nozle according to claim 2 wherein said middle lands are
disposed closer to said vane trailing edges than said leading edges
along said pressure sides, and closer to said vane leading edges
than said trailing edges along said suction sides.
5. A nozzle according to claim 4 wherein said aft-step forward
lands, middle lands, and forward-step aft lands are disposed in
said inner bands.
6. A nozzle according to claim 4 wherein said aft-step forward
lands, middle lands, and forward-step aft lands are disposed in
said outer bands.
7. A nozzle according to claim 4 wherein said aft-step forward
lands, middle lands, and forward-step aft lands are disposed in
both said outer and inner bands.
8. A nozzle according to claim 4 wherein said bands at said middle
lands have different thicknesses to effect said nominally flush
alignment therebetween.
9. A nozzle according to claim 4 wherein said middle lands have a
smaller manufacturing tolerance in radial position than said
forward and aft lands at said splitlines.
10. A method of making said nozzle according to claim 1 comprising:
casting separately said vanes and bands; bonding together said cast
vanes and bands in a pair of said nozzle segments; assembling
together said pair of nozzle segments; measuring any step between
said middle lands; and re-casting another pair of said nozzle
segments to reduce said measured step to said nominal flush
alignment at said middle lands.
11. A method according to claim 10 further comprising: casting said
vanes and bands from corresponding dies therefor; locally polishing
one of said band dies to remove material therefrom for locally
increasing thickness of a corresponding one of said middle lands to
reduce said measured step; and re-casting said segments using said
polished band die to reduce said measured step.
12. A method according to claim 11 further comprising: fixturing
said cast outer and inner bands at three points each on flow
surfaces thereof; and bonding together said vanes in said fixtured
bands.
13. A method according to claim 12 wherein said bands are fixtured
at three points corresponding with said forward and aft lands at
one end thereof and said forward land at an opposite end
thereof.
14. A turbine nozle comprising: a row of nozzle segments each
including a pair of vanes mounted at opposite ends to outer and
inner bands; each of said vanes having axially opposite leading and
trailing edges and circumferentially opposite pressure and suction
sides extending radially between said opposite ends; each of said
bands having circumferentially opposite first and second ends
adjoining at respective splitlines therebetween; each of said band
ends having a forward land extending along said splitline near said
vane leading edge, an aft land extending along said splitline near
said vane trailing edge, and a middle land extending along said
splitline over an intermediate portion of said vane between said
forward and aft lands to collectively form flowpath surfaces
bounding combustion gas flow between said vanes; said forward lands
at said splitlines having a nominal aft-facing step, said aft lands
at said splitlines having a nominal forward-facing step and said
middle lands being nominal flush; and said middle lands have a
smaller manufacturing tolerance in radial position than said
forward and aft lands at said splitlines.
15. A nozzle according to claim 14 wherein said bands at said
middle lands have different thicknesses to effect said nominally
flush alignment therebetween.
16. A nozzle according to claim 15 wherein said forward and aft
lands blend with said middle lands for blending said aft-facing
steps and forward-facing steps with said nominally flush middle
lands.
17. A nozzle according to claim 16 wherein said middle lands are
disposed closer to said vane trailing edges than said leading edges
along said pressure sides, and closer to said vane leading edges
than said trailing edges along said suction sides.
18. A nozzle according to claim 17 wherein said aft-step forward
lands, middle lands, and forward-step aft lands are disposed in
both said outer and inner bands.
19. A nozzle according to claim 18 wherein said middle lands are
sized in length along said splitlines to prevent streamlines of
combustion gases from flowing into said aft-facing steps at said
forward lands, and from flowing into said forward-facing steps at
said aft lands.
20. A nozzle according to claim 19 wherein said splitlines are
disposed obliquely between said segments, and said forward lands
are longer than each of said middle and aft lands along said
splitlines.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates generally to gas turbine
engines, and, more specifically, to turbine nozzles therein.
[0002] In a gas turbine engine air is pressurized in a compressor
and mixed with fuel in a combustor for generating hot combustion
gases that flow downstream through a high pressure turbine nozzle
which directs the flow into a row of high pressure turbine rotor
blades. The blades extract energy from the gases for powering the
compressor, and a low pressure turbine is disposed downstream
therefrom for extracting additional energy which typically powers a
fan for producing propulsion thrust to power an aircraft in
flight.
[0003] The high pressure turbine nozzle receives the highest
temperature combustion gases directly from the combustor and is
specifically configured for withstanding those gases for a useful
service life. The nozzle is an annular structure but is segmented
into arcuate segments to accommodate the substantial expansion and
contraction of the components thereof due to the hot operating
environment. Each segment includes arcuate outer and inner bands
supporting a pair of hollow stator vanes which receive a portion of
pressurized air bled from the compressor for cooling the nozzle
segments during operation.
[0004] The two bands define the radially outer and inner flowpath
surfaces between which the combustion gases are confined during
operation. The bands are separated from each by corresponding axial
splitlines which are suitably sealed with typical spline seals
therebetween.
[0005] The nozzle vanes have a crescent profile with substantial
curvature or camber between the leading and trailing edges thereof,
with a generally concave pressure side and a generally convex
opposite suction side along which the combustion gases flow during
operation. The suction side of one vane is circumferentially spaced
from the pressure side of an adjacent vane to define a flow channel
therebetween for the combustion gases. The combustion gases enter
these flow channels in a general axial downstream direction and are
redirected at an oblique angle from the outlet of the channels
defined between adjacent vane trailing edges.
[0006] Accordingly, the individual streamlines of the combustion
gases flow generally parallel to each other between the nozzle
vanes, but vary in curvature to correspond with the different
velocities thereof as effected by the suction and pressure sides of
adjacent vanes.
[0007] The band splitlines are straight and oriented obliquely in
the bands between the corresponding arcuate profiles of the
adjacent vane suction and pressure sides. Accordingly, the
combustion gases typically cross the splitline twice during their
passage between the vanes as they curve between the suction and
pressure sides and flow axially aft along the splitlines.
[0008] The bands are circumferentially continuous between each pair
of vanes in each nozzle segment and enjoy maximum aerodynamic
efficiency. However, the splitlines between the vanes of adjacent
nozzle segments provide a local discontinuity in the bands which
can affect aerodynamic efficiency.
[0009] The band flow surfaces are designed to be substantially
flush with each other at the splitlines, but due to normal
manufacturing tolerances and stack-up of those tolerances during
assembly of the nozzle components, differences in radial elevation
of the adjoining bands randomly occur with corresponding steps in
the flow path surfaces. If the step faces forwardly opposite to the
direction of the combustion gases, they introduce a local obstacle
to the smooth flow of those gases which both reduces aerodynamic
efficiency of the nozzle and locally heats the exposed edges
leading to oxidation thereof over time. Oxidation of the exposed
splitline edges reduces the useful life of the nozzle segments and
requires earlier replacement thereof than would be otherwise
required.
[0010] Since the combustion gases typically cross the oblique
splitlines twice as they pass through the nozzle channels, the
undesirable flow-obstructing steps may occur at either the forward
portion of the bands or the aft portion of the bands or may vary
therebetween in a transition zone therebetween. Since a typical
aircraft gas turbine engine operates over different power levels
from idle to maximum power, the configuration of the streamlines
through the nozzle correspondingly varies.
[0011] Accordingly, a downstream facing step during some operation
of the engine may change to an upstream facing step as the
configuration of the streamlines changes. Undesirable oxidation of
the upstream facing edge remains a practical problem due to
real-life manufacturing tolerances and the inability to exactly
dimension the nozzle components. Nozzle durability is therefore
affected by the exposed splitline edges which shortens the useful
life of the nozzle in practice.
[0012] It is, therefore, desired to provide an improved turbine
nozzle including an improved splitline configuration for enhancing
durability and useful life of the nozzle.
BRIEF SUMMARY OF THE INVENTION
[0013] A turbine nozzle includes segments of outer and inner bands
supporting corresponding vane pairs. The bands adjoin each other at
corresponding ends along splitlines, with each band having a
forward land, an opposite aft land, and a middle land extending
therebetween. The forward lands have a nominal aft-facing step, the
aft lands have a nominal forward-facing step and the middle lands
are nominally flush.
BRIEF DESCRIPTION OF THE DRAWINGS
[0014] The invention, in accordance with preferred and exemplary
embodiments, together with further objects and advantages thereof,
is more particularly described in the following detailed
description taken in conjunction with the accompanying drawings in
which:
[0015] FIG. 1 is an axial sectional view of a portion of a turbofan
aircraft gas turbine engine including a combustor discharging
combustion gases to a high pressure turbine nozzle in accordance
with an exemplary embodiment of the present invention.
[0016] FIG. 2 is an isometric view of a portion of the turbine
nozzle illustrated in FIG. 1.
[0017] FIG. 3 is a radial section view of a portion of adjoining
nozzle segments illustrated in FIG. 2 and taken along line 3-3.
[0018] FIG. 4 is a sectional view of a portion of adjoining inner
bands of the nozzle segments illustrated in FIG. 3 and taken
generally along line 4-4.
[0019] FIG. 5 is an exploded view of an exemplary one of the nozzle
segments illustrated in FIG. 2 in conjunction with a flowchart
representation of the manufacture thereof in accordance with an
exemplary embodiment of the present invention.
[0020] FIG. 6 is a radial section view through the adjoining nozzle
segments illustrated in FIG. 3 and taken generally along line
6-6.
DETAILED DESCRIPTION OF THE INVENTION
[0021] Illustrated schematically in FIG. 1 is a gas turbine engine
10 in the exemplary form of a turbofan aircraft engine configured
for powering an aircraft in flight. The engine includes in serial
flow communication a fan 12, a multistage axial compressor 14, an
annular combustor 16, a high pressure turbine nozzle 18, a high
pressure turbine 20, and a low pressure turbine 22.
[0022] During operation, air is channeled through the fan and
compressor and mixed with fuel in the combustor for generating hot
combustion gases 24 which are discharged downstream through the
high and low pressure turbine components. The high pressure turbine
20 includes a row of rotor blades which extract energy from the
gases for powering the compressor in a conventional manner.
Additional energy is extracted from the combustion gases in the
rotor blades of the low pressure turbine which in turn powers the
fan in a conventional manner.
[0023] The high pressure turbine nozzle 18 illustrated in FIG. 1 is
an annular assembly of parts fixedly mounted coaxially in the
engine on an annular support ring 26 specifically configured
therefor, which is suitably attached to an inner casing of the
engine. The nozzle includes a row of circumferentially adjacent
nozzle segments 18a, illustrated in FIG. 2, which interrupt the
circumferential continuity of the nozzle for reducing thermal
stresses therein under expansion and contraction caused by the hot
combustion gases.
[0024] As shown in FIG. 2, each of the nozzle segments includes a
pair of stator vanes 28 fixedly mounted at opposite radial ends to
corresponding arcuate outer and inner bands 30,32. Each of the
vanes includes axially opposite leading and trailing edges 34,36
and circumferentially opposite pressure and suction sides 38,40
extending radially between opposite ends of the vanes which are
fixedly mounted in corresponding apertures in the bands.
[0025] The vane pressure side 38 is generally concave, and the
suction side is generally convex for channeling the combustion
gases in the flow channels between adjacent vanes in a conventional
manner for discharge into the turbine blades. The nozzle vanes are
hollow and include various rows of film cooling holes through the
walls thereof through which a portion of air bled from the
compressor is channeled for cooling the nozzle segments during
operation in a conventional manner.
[0026] As shown in FIGS. 2 and 3, each of the inner bands has
circumferentially opposite first and second ends 42,44 which adjoin
each other at respective axially extending splitlines 46. The inner
band defines the inner flowpath boundary or surface of each nozzle
segment with the outer band correspondingly defining the outer
flowpath surface. Accordingly, the outer band also includes
circumferentially opposite first and second ends like those in the
inner band, and a corresponding splitline between adjacent ones of
the outer bands.
[0027] As further described hereinbelow, the inner bands 32 are
specifically configured at the splitlines for improving aerodynamic
performance and reducing undesirable oxidation thereat. The outer
bands may be similarly configured in the same manner at the
splitlines therebetween as further described hereinbelow.
[0028] More specifically, the splitlines 46 in the outer and inner
bands are required for reducing thermal stress in the nozzle during
operation, but introduce discontinuities along the circumference of
the nozzle. The band ends face each other at the corresponding
splitlines and include hidden slots therein in which are disposed
conventional spline seals that seal the adjoining ends of the
nozzle segments in a conventional manner.
[0029] As indicated above, due to manufacturing tolerances and
stack-up of those tolerances during the assembly of the nozzle
components, the junction of the nozzle bands at the corresponding
splitlines can randomly have undesirable upstream facing steps
which locally block the downstream flow of the combustion gases and
is subject to local heating and oxidation which would decrease
nozzle efficiency and useful life in service.
[0030] However, the band joints at the corresponding splitlines are
specifically configured in accordance with the present invention
for reducing or eliminating the random effect of upstream facing
steps for improving efficiency of the nozzle, decreasing oxidation
at the splitlines, and increasing useful service life.
[0031] More specifically, and initially referring to FIG. 3, each
of the band ends has a forward land 42a,44a extending aft from the
front edge of the band along the splitline near the vane leading
edge, and an aft land 42b,44b extending forward from the aft end of
the band along the splitline near the vane trailing edge. And, each
band end also includes a middle land 42c,44c extending along the
splitline over an intermediate portion of the vane between the
respective forward and aft lands to collectively form flowpath
surfaces which radially bound the combustion gas flow between the
vanes.
[0032] In accordance with the present invention, the forward lands
42a,44a adjoining at the respective splitlines 46 have a nominal
aft-facing step 48, and the aft lands 42b,44b adjoining at the same
splitlines have a nominal forward-facing step 50. And, most
significantly, the middle lands 42c,44c adjoining at the same
splitlines are nominally flush with each other, with no step
nominally therebetween.
[0033] As indicated above, the manufacture of turbine components is
subject to random variation in all dimensions, since perfect or
exact dimensions are impossible to achieve. As indicated below, the
various components of the turbine nozzle are initially cast,
assembled together, and brazed in two-vane nozzle segments which
are assembled together in the annular nozzle itself supported by
axial bolts through the annular support 26 illustrated in FIG.
1.
[0034] The final radial position of the flowpath surfaces of the
inner and outer bands is determined by the manufacturing tolerances
of the bands themselves, the assembly tolerances when joined to the
vanes, and the additional assembly tolerances when mounted to the
annular support 26. These various tolerances randomly add together
in the finally assembled turbine nozzle, with the inner and outer
bands being radially aligned with each other within an acceptable
dimensional tolerance range suitably greater than and less than a
nominal value representing a desired final dimensional position of
the band flowpath surfaces.
[0035] In the fabrication and assembly of conventional turbine
nozle components, the radial position of the flowpath surfaces of
the inner and outer bands at the splitlines may be obtained within
a dimensional tolerance of about plus or minus twenty (+/-20) mils.
For a conventional turbine nozzle in which the band ends are
designed for a nominally flush final position, random variation in
the dimensional tolerance can lead to substantial forward or aft
facing steps which locally obstruct the combustion gas flow, are
locally heated thereby, and are subject to oxidation reducing the
useful life of the turbine nozzle.
[0036] However, and in accordance with the present invention, the
selectively introduced aft-facing steps 48 at the forward lands
ensure downstream flow of the combustion gases without impingement
against an outwardly protruding band edge. This is illustrated in
FIGS. 3 and 4 in which the downstream flowing combustion gases 24
flow over the forward lands 44a of the left illustrated band and
drop to the relatively lower forward land 42a on the right
band.
[0037] Similarly, the forward-facing steps 50 at the aft lands
42b,44b ensure that as the combustion gases change direction
between the adjacent vanes over the splitline, they again drop from
the aft land 42b of the right band to the lower aft land 42b on the
left band illustrated.
[0038] The middle lands 42c,44c represent a transition zone between
the forward and aft lands in which the streamlines of the
combustion gas flow generally parallel with the splitline and
undergo a change in direction between the adjacent bands.
[0039] As best shown in FIG. 3, the splitlines 46 are disposed
obliquely between the adjacent nozzle segments, with the
corresponding band ends 42,44 also being oblique for positioning
the splitline generally uniformly between the obliquely oriented
nozzle vanes. This oblique orientation is conventional for
redirecting the combustion gases which enter the nozzle flowpath
channels generally axially, and discharging the combustion gases in
a substantial tangential direction for engaging the downstream
rotor blades. In this configuration, the forward lands 42a,44a
illustrated in FIG. 3 are longer along the splitline than each of
the middle and aft lands 42c,44c and 42b,44b.
[0040] With this oblique configuration of the splitlines 46, the
combustion gases must necessarily flow over the forward portion
thereof from one band to the next, and then again flow in an
opposite direction over the splitline back to the initial band as
the combustion gases are discharged from the nozzle outlet.
Accordingly, the combustion gas streamlines transition between
these two directions, which transition preferably occurs in the
middle lands 42c,44c.
[0041] The middle lands are therefore predeterminedly sized in
length A along a limited portion of the splitlines to preferably
prevent streamlines of the combustion gases from flowing into the
aft-facing steps 48 at the forward lands, and from flowing into the
forward-facing steps 50 over a preferred operating range of the
engine, such as cruise power and extending down to idle or up to
maximum power as desired. The middle lands therefore have a finite
length A and are not merely a singular point in the transition
between the forward and aft facing steps to ensure that the
combustion gases do not flow upstream over the deliberately
introduced forward and aft facing steps.
[0042] For example, if the middle lands were eliminated, with the
aft and forward lands joining each other near the middle of the
splitlines, it would be possible for some of the streamlines
flowing along the right band past the vane leading edges
illustrated in FIG. 3 to engage the aft-facing steps as they turn
early over the splitlines. Alternatively, it would also be possible
for some of the streamlines from the left band to engage a portion
of the forward-facing step due to late turning of the streamlines
over the splitlines.
[0043] The introduction of the middle lands being substantially
flush with each other reduces the likelihood that early or late
flow turning between the vanes will flow upstream over the forward
and aft facing steps instead of downstream thereover.
[0044] In the preferred embodiment illustrated in FIGS. 3 and 4,
the forward and aft lands preferably blend smoothly with the
corresponding middle lands for blending both the aft-facing steps
48 and the forward-facing steps 50 with the nominally flush or no
step portion of the splitlines effected by the middle lands.
[0045] As shown schematically in FIG. 4, the forward land 44a of
the left band is disposed at a higher elevation (+) relative to the
forward land 42a of the right band (-) with the difference in
elevation thereof representing the height B of the forward-facing
step 48. Similarly, the aft land 42b of the right band is at a
higher elevation (+) relative to the aft land 44b of the left band
(-) as represented by the difference in relative elevation defining
the height C of the forward-facing step 50.
[0046] The two step heights B,C preferably decrease in magnitude
from the opposite forward and aft ends of the bands toward the
middle lands therebetween with which the forward and aft lands
blend to a substantially zero difference in elevation for providing
the nominally flush alignment therebetween.
[0047] In the exemplary embodiment illustrated in FIG. 3, the
middle lands 42c,44c are disposed closer to the vane trailing edges
than the leading edges along the pressure sides 38 of the vanes.
And, the middle lands are also disposed closer to the vane leading
edges than the trailing edges along the corresponding suction sides
40 of the vanes. In view of the different aerodynamic effects of
the pressure and suction sides of the vanes, the transition region
in which the combustion gas streamlines change direction for
recrossing the splitlines occurs near the vane trailing edge on the
pressure side of one vane and near the intermediate hump region on
the suction side of the opposite vane. The middle lands are
therefore preferably located and sized for covering this transition
region for any suitable operating range of the engine.
[0048] The introduction of the forward lands and aft-facing steps
48, flush middle lands, and aft lands with forward-facing steps 50
is preferred in the inner bands 32, or in the outer bands 30, and
is preferably found in both the outer and inner bands.
[0049] In view of the random nature of final position of the
flowpath surfaces at the splitlines due to manufacturing tolerances
and assembly stack-up thereof, the introduction of the three
discrete forward, aft, and middle lands along the splitlines may be
used to advantage for decreasing the likelihood and extent of a
splitline edge protruding into a portion of the combustion gas flow
which would locally increase the temperature thereof and lead to
excessive oxidation reducing nozzle life.
[0050] By deliberately and selectively introducing the aft-facing
and forward-facing steps 48,50 with the nominally flush middle
lands therebetween, oxidation of the splitline edges may be reduced
for increasing nozzle durability and service life. This advantage
may be achieved with normal manufacturing tolerances.
[0051] However, in accordance with another feature of the present
invention, the middle lands 42c,44c are preferably manufactured
with a smaller manufacturing tolerance in radial position of the
flowpath surfaces thereof than the manufacturing tolerance for the
forward and aft lands along the splitline. In this way, improved
flushness of the adjoining middle lands may be effected for further
reducing the likelihood and magnitude of any step introduced
therebetween in practice.
[0052] FIG. 5 illustrates schematically a preferred method of
making the individual nozzle segments which collectively form the
annular turbine nozzle. The individual nozzle vanes 28 and outer
and inner bands 30,32 are separately cast using master or metal
dies or molds 28M,30M, and 32M. These molding dies are used in a
conventional manner of casting using the lost wax method.
[0053] In this method, the molding dies are complementary with the
external surfaces of the vanes and bands and are initially filled
with wax. The wax is solidified, removed from the dies, and then
coated with ceramic to form a molding shell. The wax is removed
from the shell and replaced by molten metal which forms the
corresponding parts. Since the vanes are preferably hollow,
conventional ceramic cores are used in conjunction with the ceramic
shells for casting the vanes in any conventional manner.
[0054] The cast vanes and bands are then assembled together by
inserting the corresponding opposite hub ends of the vanes into
corresponding seats in the bands, and temporarily held together in
suitable fixtures therefor.
[0055] The vane hubs are then brazed into their corresponding seats
in the bands for fixedly bonding together the two cast vanes and
two bands in a four-part assembly for each nozzle segment. Each
nozzle segment is then further processed by drilling the various
rows of film cooling holes in the vanes and other required features
for completing a full row of nozzle segments defining the completed
turbine nozzle.
[0056] In a preferred embodiment, a pair of the nozzle segments are
manufactured in this manner using the same master dies for the
vanes and bands, and then the two nozzle segments are assembled
together in a suitable fixture, or on the intended annular support
26 illustrated in FIG. 1, for achieving the intended relative
position thereof in the turbine nozzle.
[0057] In conventional casting of turbine nozzle parts, the master
dies are configured for the nominal dimensions of the corresponding
vane and band parts, and are therefore subject to random variations
in dimensions, which are substantial in cast parts as compared with
machined parts which may have closer or smaller variation in
tolerances. The nominal casting of the nozle parts and stack-up
tolerances thereof during assembly will typically result in a
random variation in relative elevation between the adjoining middle
lands 44c as illustrated schematically in FIG. 6.
[0058] As a result thereof, the desired flushness of the adjacent
middle lands may vary within a corresponding tolerance range, with
the initial introduction of a locally small middle step having a
magnitude D. That middle step D is measured while the two adjacent
nozzle segments are suitably fixtured in relative position. And
then, another pair of the nozzle segments are preferably re-cast to
reduce the measured step D to achieve the desired nominal flush
alignment of the middle lands.
[0059] More specifically, the relevant portion of the molding die
32M for the inner band 32 is illustrated in FIG. 6 which is
preferably locally polished to remove material therefrom which will
correspondingly locally increase thickness of the corresponding
middle land 44c when the next band is cast using the polished die.
In this way, the initial step D measured between the two middle
lands 44c may be substantially reduced or eliminated in subsequent
castings by correspondingly polishing the respective portion of the
molding die therefor.
[0060] The casting of sample nozzle segments, assembly thereof,
measuring of any difference in elevation between the middle lands,
locally polishing the band dies, and re-casting a new set of nozzle
segments may be done one or more times as desired for reducing or
substantially eliminating the magnitude of the initial middle step
D. The resulting inner bands will therefore have different
thicknesses at the splitline to effect the nominal flush alignment
therebetween. As shown in FIG. 6, a portion of the left middle land
44c is illustrated in phantom line with its initial step D, which
step is substantially eliminated as shown in solid line by the
local polishing of the corresponding die 32M which will locally
introduce additional material in the cast band.
[0061] Since the splitline 46 is defined between opposite portions
of the middle lands 44c on adjacent inner bands, one of those
portions will remain with its original thickness as initially cast,
with the other portion being locally thickened to substantially
eliminate any surface step therebetween. In this way, the relative
degree of flushness between the middle lands of adjoining bands may
be substantially improved over conventional casting in which the
dies are not so polished. For example, the conventional
manufacturing tolerance of about +/-20 mils on the radial position
of the band flowpath surfaces may be substantially reduced to about
plus or minus six (+/-6) mils in the method described above.
[0062] As indicated above the three-land configuration along the
splitlines is preferably introduced for both the inner and outer
bands, and correspondingly the molding dies therefor may be
similarly locally polished as required for achieving the desired
flushness in the middle lands between the adjoining inner and outer
bands at which the corresponding aft and forward facing steps
blend.
[0063] As illustrated schematically in FIG. 5, the cast outer and
inner bands 30,32 are preferably fixtured or physically grounded at
three reference points each on the corresponding flow surfaces
thereof. This three-point fixturing ensures accurate alignment of
the bands so that the vanes 28 may be accurately brazed in the
fixtured bands.
[0064] In the preferred embodiment illustrated in FIG. 5, the
corresponding outer and inner bands 30,32 are fixtured at three
points corresponding with the forward and aft lands 42a,b at one
end thereof and the forward land 44a at an opposite end thereof. In
this way the relative radial position of the flowpath surfaces at
the adjoining splitlines may be more accurately achieved for the
nominally flush middle lands and stepped forward and aft lands.
[0065] Furthermore, with the brazed nozzle segments being so
fixtured at three points, two corresponding mounting holes 52 may
be accurately drilled in the radial support flange extending
inwardly from the inner band. In this way, when the nozzle segments
are bolted to the annular support 26 illustrated in FIG. 1, precise
alignment of the adjoining inner bands and the adjoining outer
bands may be effected with the desired stepped forward and aft
lands and nominally flush middle lands therebetween.
[0066] Notwithstanding conventional manufacturing tolerances and
buildup thereof as described above, the adjoining inner and outer
bands of the nozzle segments may be preferentially configured with
the aft-facing steps at the forward lands, forward-facing steps at
the aft lands, and the substantially flush middle lands
therebetween for maximizing aerodynamic efficiency of the turbine
nozzle without locally obstructing flow of the combustion gases due
to protruding splitline edges. The splitline edges are protected
from the combustion gases and are not additionally heated thereby
and oxidized as they would if protruding into the combustion gas
flow. Durability of the turbine nozzle is therefore improved for
correspondingly improving the useful life thereof.
[0067] 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.
* * * * *