U.S. patent number RE39,479 [Application Number 10/358,927] was granted by the patent office on 2007-01-23 for durable turbine nozzle.
This patent grant is currently assigned to General Electric Company. Invention is credited to Glenn H. Nichols, Judd Dodge Tressler.
United States Patent |
RE39,479 |
Tressler , et al. |
January 23, 2007 |
Durable turbine nozzle
Abstract
A turbine nozzle includes a plurality of vanes joined at
opposite ends to outer and inner bands. The inner band has a
forward hook which is segmented to reduce thermal mismatch. And, in
additional embodiments the vane includes an impingement baffle
having preferential cooling.
Inventors: |
Tressler; Judd Dodge (Mason,
OH), Nichols; Glenn H. (Mason, OH) |
Assignee: |
General Electric Company
(Schenectady, NY)
|
Family
ID: |
23046973 |
Appl.
No.: |
10/358,927 |
Filed: |
February 5, 2003 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
|
Reissue of: |
09274144 |
Mar 22, 1999 |
06183192 |
Feb 6, 2001 |
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Current U.S.
Class: |
415/115;
415/209.2; 415/191; 415/209.3; 416/95; 415/211.2; 415/116 |
Current CPC
Class: |
F01D
9/042 (20130101); F01D 5/189 (20130101); Y02T
50/673 (20130101); Y02T 50/60 (20130101); Y02T
50/67 (20130101); Y02T 50/671 (20130101); Y02T
50/676 (20130101) |
Current International
Class: |
F01D
9/04 (20060101) |
Field of
Search: |
;415/115,191,208.2,209.2,209.3,209.4,211.2,116
;416/96A,96R,95,97R |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Nguyen; Ninh H.
Attorney, Agent or Firm: Andes; William Scott Armstrong
Teasdale LLP
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 for a gas turbine engine, comprising: a
plurality of vanes integrally joined at opposite ends to outer and
inner bands; and said inner band having a forward hook inboard of a
leading edge thereof, said hook including an inner lip and a
plurality of circumferentially spaced apart outer lips spaced
radially from said inner lip to define a retention slot
therein.
2. A nozzle according to claim 1 wherein said inner band is an
arcuate segment, and said outer lips are disposed at opposite
circumferential ends thereof and intermediate therebetween.
3. A nozzle according to claim 1 wherein said vanes join said outer
band at a compound radii fillet.
4. A nozzle according to claim 3 wherein said fillet has a larger
radius adjacent said vane than adjacent said outer bands.
5. A nozzle according to claim 1 wherein: said vanes have a
generally concave, pressure side and an opposite, generally convex,
suction side extending between leading and trailing edges and
between said outer and inner bands along respective spans of said
vanes; and said vanes include cavities having an impingement
baffle, and said baffles have a plurality of spaced apart
impingement holes extending therethrough and arranged in different
patterns facing an inner surface of said cavities for
preferentially channeling cooling air in impingement jets toward
midspan of said vanes.
6. A nozzle according to claim 5 wherein said impingement hole
pattern on said vane pressure side is denser than said pattern on
said vane suction side for preferentially cooling said vane
midspan.
7. A nozzle according to claim 6 wherein said baffles include
imperforate regions at opposite span ends thereof that converge
toward said midspan between said leading and trailing edges
thereof.
8. A nozzle according to claim 7 wherein: said impingement holes on
said pressure and suction sides have equal size; and said baffles
include a row of larger impingement holes extending along leading
edges thereof.
9. A nozzle according to claim 5 wherein said baffles include a row
of larger impingement holes extending along leading edges
thereof.
10. A nozzle according to claim 9 wherein said impingement holes
have a greater flow density inside said vane leading edge than on
said vane pressure and suction sides.
11. A nozzle according to claim 9 wherein said baffle leading edge
has a radius sufficient for forming said large impingement holes
planar therein, and said vane leading edge has a correspondingly
larger radius to complement said baffle leading edge.
12. A nozzle according to claim 5 wherein said baffles complement
said vane cavities for maintaining a uniform gap with said inner
surfaces thereof between said vane leading and trailing edges.
13. A nozzle according to claim 5 wherein said baffles include
integral standoff pads for spacing said baffles from said vane
inner surfaces, and said pads are arranged on opposite sides of
said baffles at both leading and trailing edges thereof.
14. A nozzle according to claim 13 wherein said pads are more
uniformly spaced on said vane pressure side said than said suction
side.
15. A turbine nozzle for a gas turbine engine, comprising: a
plurality of vanes integrally joined at opposite ends to outer and
inner bands; said inner band having a forward hook inboard of a
leading edge thereof, said hook including an inner lip and a
plurality of circumferentially spaced apart outer lips spaced
radially from said inner lip to define a retention slot therein;
said vanes have generally concave, pressure sides and opposite,
generally convex, suction sides extending between leading and
trailing edges thereof and between said outer and inner bands along
respective spans of said vanes; and said vanes include cavities
each having an impingement baffle, and said baffles have a
plurality of spaced apart impingement holes extending therethrough
and arranged in different patterns facing an inner surface of said
cavities for preferentially channeling cooling air in impingement
jets toward midspan of said vane.
16. A nozzle according to claim 15 wherein said baffles include
imperforate regions at opposite span ends thereof that converge
toward said midspan between said leading and trailing edges
thereof.
17. A nozzle according to claim 16 wherein said baffles include a
row of larger impingement holes extending along leading edges
thereof.
18. A nozzle according to claim 17 wherein said inner band is an
arcuate segment, and said outer lips are disposed at opposite
circumferential ends thereof and intermediate therebetween.
19. A nozzle according to claim 18 wherein said vanes join said
outer band at a compound radii fillet.
.Iadd.20. A turbine nozzle for a gas turbine engine, said nozzle
comprising: a radially inner band; a radially outer band; and at
least one vane extending between said inner and outer bands, said
vane comprising a first sidewall and a second sidewall connected
together at a leading edge and a trailing edge to define a cavity
therebetween; and an impingement baffle positioned within the vane
cavity, said impingement baffle comprising a plurality of first
impingement baffle holes extending therethrough and a plurality of
second impingement baffle holes extending therethrough, at least
some of said first impingement baffle holes extending along said
vane leading edge and having a first diameter, said second
impingement baffle holes having a second diameter, said first
diameter is larger than said second diameter, said impingement
baffle leading edge having a radius sufficient for forming said
first impingement baffle holes substantially planar therein, said
at least one vane leading edge having a corresponding larger radius
that substantially compliments said impingement baffle leading
edge..Iaddend.
.Iadd.21. A turbine nozzle in accordance with claim 20 wherein said
plurality of first impingement baffle holes for channeling cooling
fluid therethrough in impingement jets towards an inner surface of
said at least one vane leading edge..Iaddend.
.Iadd.22. A turbine nozzle in accordance with claim 20 wherein said
first sidewall defines a pressure side of said at least one vane,
said second sidewall defines a suction side of said at least one
vane, said plurality of second impingement baffle holes arranged in
a denser pattern along said pressure side than said suction
side..Iaddend.
.Iadd.23. A turbine nozzle in accordance with claim 20 wherein said
plurality of second impingement baffle holes are each sized
identically..Iaddend.
.Iadd.24. A turbine nozzle in accordance with claim 20 wherein said
plurality of first impingement baffle holes comprises a row of said
first impingement holes each having said first diameter larger than
said second diameter of each of said plurality of second
impingement baffle holes..Iaddend.
.Iadd.25. A turbine nozzle in accordance with claim 24 wherein said
row of first impingement holes extends at least partially between
said radially inner and outer bands..Iaddend.
.Iadd.26. A turbine nozzle in accordance with claim 20 wherein at
least one of said inner band and said outer band comprises a
forward hook inboard of a leading edge thereof, said hook
comprising an inner lip and a plurality of circumferentially-spaced
apart outer lips spaced radially from said inner lip such that a
slot is defined therein..Iaddend.
.Iadd.27. A turbine nozzle in accordance with claim 20 further
comprising a compound radii fillet extending between said at least
one vane and at least one of said inner band and said outer
band..Iaddend.
.Iadd.28. A turbine nozzle in accordance with claim 27 wherein said
compound radii fillet comprises a first radius and a second radius,
said second radius larger than said first radius, said first radius
between said second radius and at least one of said inner band and
said outer band..Iaddend.
.Iadd.29. A turbine nozzle in accordance with claim 20 wherein said
at least one vane further comprises a plurality of projections
extending outwardly from said impingement baffle, said projections
configured to maintain a relative position of said impingement
baffle with respect to said at least one vane such that a
substantially uniform gap is defined between said impingement
baffle and said at least one vane..Iaddend.
.Iadd.30. A turbine nozzle for a gas turbine engine, said nozzle
comprising: a radially inner band; a radially outer band; and at
least one vane extending between said inner and outer bands, said
vane comprising a first sidewall and a second sidewall connected
together at a leading edge and a trailing edge to define a cavity
therebetween, said at least one vane formed integrally with said
radially outer band such that a compound fillet extends between
said radially outer band and said at least one vane, said compound
fillet circumscribes said at least one vane, and comprises a first
radius and a second radius that is larger than said first radius,
said first radius extending between said second radius and at least
one of said radially outer and inner bands..Iaddend.
.Iadd.31. A turbine nozzle in accordance with claim 30 wherein said
first sidewall defines a pressure side of said at least one vane,
said second sidewall defines a suction side of said at least one
vane, each said sidewall comprises a plurality of impingement
baffle holes extending therethrough, said plurality of impingement
baffle holes are arranged in a denser pattern along said vane
pressure side than along said vane suction side..Iaddend.
.Iadd.32. A turbine nozzle in accordance with claim 30 wherein at
least one of said first and said second sidewall comprises a
plurality of impingement baffle holes extending therethrough, said
plurality of impingement baffle holes comprise a first row of
impingement holes and a pattern of remaining impingement holes, at
least some of said first row of impingement holes having a first
diameter, said pattern of remaining impingement holes each having a
second diameter that is smaller than said first diameter, said
first row of impingement baffle holes extending along said vane
leading edge..Iaddend.
.Iadd.33. A turbine nozzle in accordance with claim 30 wherein said
inner band comprises a forward hook inboard of a leading edge
thereof, said hook comprising an inner lip and a plurality of
circumferentially-spaced apart outer lips spaced radially from said
inner lip such that a slot is defined therein..Iaddend.
.Iadd.34. A turbine nozzle for a gas turbine engine, said nozzle
comprising at least one vane comprising a first sidewall and a
second sidewall connected together at a leading edge and a trailing
edge to define a cavity therebetween, the cavity sized to receive
an impingement baffle therein, said impingement baffle comprising a
plurality of impingement holes extending therethrough, said
plurality of impingement holes comprising at least a row of first
impingement holes for directing cooling airflow along an inner
surface of said at least one vane leading edge, and a pattern of
remaining impingement holes, at least some of said impingement
holes have a first diameter, said remaining impingement holes have
a second diameter, said first diameter is different than said
second diameter, said impingement baffle leading edge has a radius
sufficient for forming said first impingement holes substantially
planar therein..Iaddend.
.Iadd.35. A turbine nozzle in accordance with claim 34 wherein said
at least one vane leading edge has a first radius that is larger
than a second radius of said impingement baffle leading edge, said
at least one vane first radius substantially compliments said
impingement baffle leading edge second radius..Iaddend.
.Iadd.36. A turbine nozzle in accordance with claim 34 wherein said
inner band comprises a forward hook inboard of a leading edge
thereof, said hook comprising an inner lip and a plurality of
circumferentially-spaced apart outer lips spaced radially from said
inner lip such that a slot is defined therein..Iaddend.
.Iadd.37. A turbine nozzle in accordance with claim 34 further
comprising a compound radii fillet extending between said outer
band and each of said plurality of vanes..Iaddend.
.Iadd.38. A turbine nozzle in accordance with claim 37 wherein each
of said compound radii fillets comprises a first radius and a
second radius, said first radius smaller than said second radius
and extending between said second radius and said outer
band..Iaddend.
.Iadd.39. A turbine nozzle in accordance with claim 34 wherein said
first diameter is larger than said second diameter..Iaddend.
.Iadd.40. A turbine nozzle in accordance with claim 34 wherein said
first sidewall defines a pressure side of said at least one vane,
said second sidewall defines a suction side of said at least one
vane, said plurality of impingement baffle holes arranged in a
denser pattern along said pressure side than said suction
side..Iaddend.
.Iadd.41. A turbine nozzle in accordance with claim 34 wherein a
gap is defined between said row of said impingement baffle leading
edge and said at least one vane leading edge, said first row of
impingement holes facilitates increasing a ratio of the cooling
area behind said at least one vane within said gap to the cooling
area external to said at least one vane leading edge..Iaddend.
.Iadd.42. A turbine nozzle in accordance with claim 41 wherein said
gap is substantially uniform between said at least one vane leading
edge and said impingement baffle leading edge..Iaddend.
.Iadd.43. An impingement baffle for a turbine nozzle for use in a
gas turbine engine, said impingement baffle comprising a first
sidewall and a second sidewall connected together at a leading edge
and a trailing edge, said impingement baffle further comprising a
plurality of rows of impingement holes extending therethrough, said
plurality of rows of impingement holes comprising at least a first
row of first impingement holes extending along said baffle leading
edge and a pattern of remaining impingement holes, at least some of
said impingement holes have a first diameter and said remaining
impingement holes have a second diameter, said first diameter is
different than said second diameter, said impingement baffle
leading edge has a radius sufficient for forming said first
diameter first impingement holes substantially planar
therein..Iaddend.
.Iadd.44. An impingement baffle in accordance with claim 43 wherein
said first diameter is larger than said second
diameter..Iaddend.
.Iadd.45. An impingement baffle in accordance with claim 43 wherein
said first sidewall defines a pressure side of said impingement
baffle, said second sidewall defines a suction side of said
impingement baffle, said plurality of impingement baffle holes
arranged in a denser pattern along said pressure side than said
suction side..Iaddend.
.Iadd.46. An impingement baffle in accordance with claim 43 further
comprising a plurality of standoff pads extending outwardly, from
at least one of said first sidewall and said second
sidewall..Iaddend.
.Iadd.47. An impingement baffle in accordance with claim 43 wherein
the turbine nozzle includes at least one vane, said first row of
first impingement holes facilitates impingement cooling of a
leading edge of the turbine nozzle vane..Iaddend.
.Iadd.48. A turbine nozzle for a gas turbine engine, said nozzle
comprising: a radially inner band; a radially outer band; and at
least one vane extending between said radially inner and outer
bands, said inner band comprising a leading edge and a forward hook
radially inward of said leading edge, said forward hook comprising
a radially inner lip and a plurality of circumferentially-spaced
apart radially outer lips such that a retention slot is defined
therebetween..Iaddend.
.Iadd.49. A turbine nozzle in accordance with claim 48 wherein said
at least one vane is formed integrally with at least one of said
radially outer band and said radially inner band such that a
compound fillet extends between said at least one vane and at least
one of said radially outer and inner bands..Iaddend.
.Iadd.50. A turbine nozzle in accordance with claim 49 wherein said
compound fillet circumscribes said at least one vane and comprises
a first radius and a second radius that is larger than said first
radius, said first radius extending between said second radius and
at least one of said radially inner and outer bands..Iaddend.
.Iadd.51. A turbine nozzle in accordance with claim 48 further
comprising an impingement baffle, said at least one vane comprises
a first sidewall and a second sidewall connected together at a
leading edge and a trailing edge to define a cavity therebetween,
the cavity sized to receive said impingement baffle therein, said
impingement baffle comprising a plurality of impingement holes
extending therethrough..Iaddend.
.Iadd.52. A turbine nozzle in accordance with claim 51 wherein said
impingement baffle plurality of impingement holes comprise at least
a row of first impingement holes for directing cooling airflow
along an inner surface of said at least one vane leading edge, and
a pattern of remaining impingement holes, at least some of said
first impingement holes have a first diameter, said pattern of
remaining impingement holes have a second diameter, said first
diameter is different than said second diameter..Iaddend.
.Iadd.53. A turbine nozzle in accordance with claim 51 wherein said
impingement baffle leading edge has a radius sufficient for forming
said first impingement holes substantially planar
therein..Iaddend.
.Iadd.54. A turbine nozzle in accordance with claim 51 wherein said
at least one vane first sidewall defines a suction side of said at
least one vane, said second sidewall defines a pressure side of
said at least one vane, said impingement baffle plurality of
impingement baffle holes arranged in a denser pattern along said
vane pressure side than along said vane suction side..Iaddend.
.Iadd.55. A turbine nozzle in accordance with claim 51 wherein said
at least one vane leading edge has a first radius that is larger
than a second radius of said impingement baffle leading edge, said
at least one vane leading edge first radius substantially
compliments said impingement baffle leading edge second
radius..Iaddend.
Description
BACKGROUND OF THE INVENTION
The present invention relates generally to gas turbine engines,
and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and
mixed with fuel and ignited in a combustor for generating hot
combustion gases. Energy is extracted from the gases in
corresponding turbine stages which power the compressor and produce
useful work, such as powering a fan in a turbofan engine for
propelling an aircraft in flight, for example.
Since the turbines are bathed in the hot combustion gases during
operation, they must be suitably cooled which is typically
accomplished by bleeding a portion of the pressurized air from the
compressor and channeling it through the turbine components.
A high pressure turbine directly receives gases from the combustor
and includes a stator nozzle and a corresponding first stage rotor
having a plurality of rotor blades extending radially outwardly
from a supporting disk. A second stage nozzle then directs the
combustion gases through a corresponding row of rotor blades
extending from another rotor disk. The second stage nozzle receives
lower temperature combustion gases than the first stage nozzle and
therefore has different cooling requirements, which are typically
effected in a different manner than that for the first stage
nozzle.
Turbine nozzles are designed for durability with extensive lives
measured in thousands of hours or thousands of cycles of operation.
Such extended life is difficult to achieve since the nozzles are
subject to various differential temperatures during operation which
create thermal loads and stress therefrom. And, temperature
distributions and heat transfer coefficients of the combustion
gases channeled through the nozzle vary significantly and increase
the complexity of providing corresponding cooling. Suitable nozzle
cooling is required to limit thermal stresses and ensure a useful
life.
A typical turbine nozzle includes a row of stator vanes joined at
radially opposite ends to corresponding outer and inner bands. The
bands are typically segmented in the circumferential direction, and
include two or more vanes in corresponding sectors. The vane
sectors permit differential movement during combustion gas
temperature changes for reducing undesirable thermal stress during
operation.
The individual vanes are hollow and typically include an
impingement baffle therein which is a perforated sheet metal sleeve
spaced from the inner surface of the vane cavity for channeling
cooling air in impingement jets there against.
This type of turbine nozzle specifically configured for a second
stage turbine has enjoyed many years of commercial service in this
country. However, these nozzles are beginning to experience
distress at high cycle operation which may require their
replacement prior to their expected useful life. Nozzle distress is
caused by locally high heat transfer coefficients in different
regions of the nozzle at which corresponding cooling is limited.
Thermal gradients lead to thermal stress, which adversely affect
the useful life of the nozzle.
Accordingly, it is desired to uncover the source of high cycle
turbine nozzle distress, and improve the nozzle design for
increasing nozzle durability and corresponding life.
BRIEF SUMMARY OF THE INVENTION
A turbine nozzle includes a plurality of vanes joined at opposite
ends to outer and inner bands. The inner band has a forward hook
which is segmented to reduce thermal mismatch. And, in additional
embodiments the vane includes an impingement baffle having
preferential cooling.
BRIEF DESCRIPTION OF THE DRAWINGS
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:
FIG. 1 is an axial, partly sectional view of a turbine region of a
gas turbine engine in accordance with an exemplary embodiment of
the present invention.
FIG. 2 is an isometric view of a portion of the second stage
turbine nozzle illustrated in FIG. 1 in accordance with a preferred
embodiment of the present invention.
FIG. 3 is an isometric view of the nozzle sector illustrated in
FIG. 2 having a segmented forward hook in its inner band in
accordance with an exemplary embodiment.
FIG. 4 is a radial sectional view through one of the nozzle vanes
and attached outer band having a compound fillet in accordance with
another embodiment of the invention.
FIG. 5 is a radial sectional view through the midspan of one of the
vanes illustrated in FIG. 2 and taken along line 5--5.
FIG. 6 is an isometric view of the convex side of an impingement
baffle contained in the second stage nozzle vanes illustrated in
FIGS. 1-5.
FIG. 7 is an isometric view of the concave side of the impingement
baffle illustrated in FIG. 6.
FIG. 8 is an isometric view facing the leading edge of the
impingement baffle illustrated in FIGS. 6 and 7.
FIG. 9 is an enlarged view of a portion of the baffle illustrated
in FIG. 6 within the circle labeled 9.
FIG. 10 is an enlarged view of a portion of the baffle illustrated
in FIG. 7 within the circle labeled 10.
FIG. 11 is an enlarged view of a portion of the baffle illustrated
in FIG. 8 within the circle labeled 11.
DETAILED DESCRIPTION OF THE INVENTION
Illustrated in FIG. 1 is a portion of an exemplary aircraft gas
turbine engine 10 which is axisymmetrical about a longitudinal or
axial centerline axis 12. The engine includes a fan and a
multistage compressor (not shown) through which air 14 is
pressurized in turn, with the fan air being used for propelling an
aircraft in flight, and the air pressurized in the compressor being
mixed with fuel and ignited in a combustor 16, only the aft portion
thereof being illustrated, for generating hot combustion gases 18
which flow downstream therefrom.
The engine includes a high pressure turbine 20 having a first stage
stator nozzle 22 followed in turn by a row of first stage turbine
rotor blades 24 extending radially outwardly from a supporting
disk. The combustion gases 18 are channeled through the nozzle
vanes 22 and blades 24 for powering the compressor in a
conventional manner.
Disposed immediately downstream from the first stage blades 24 is a
second stage turbine stator or nozzle 26 which in turn channels the
combustion gases to a downstream row of second stage turbine rotor
blades (not shown) which power the fan in a conventional
manner.
But for the improved second stage turbine nozzle 26 illustrated in
FIG. 1, the engine 10 is otherwise conventional and is
representative of the engine addressed above which has enjoyed many
years of commercial use in this country. However, the originally
provided second stage nozzle therein has shown signs of distress
after a substantial number of hours and cycles of operation short
of its expected life. In accordance with the present invention,
various sources of that distress have been discovered, and the
nozzle being improved for further increasing the useful life
thereof.
More specifically, the second stage turbine nozzle 26 illustrated
in FIG. 1 includes a plurality of hollow vanes 28 circumferentially
spaced apart from each other and integrally joined at opposite
radial ends to corresponding outer and inner bands 30,32. The outer
band 30 has forward and aft hooks which support the nozzle from a
surrounding annular casing 34 in a conventional manner.
The inner band 32 is relatively thin and is locally enlarged at its
forward and aft ends for integrally including a forward hook 36 and
an aft flange 38 which extend radially inwardly for supporting a
honeycomb rotor seal 40 in a conventional manner.
As illustrated in more particularly in FIG. 2, the nozzle 26 is
preferably formed in a plurality of arcuate sectors in which the
outer and inner bands 30,32 are arcuate segments integrally formed
with two or more vanes 28 in a common casting. The forward hook 36
of the inner band 32 is disposed radially inboard of the leading
edge of the inner band and slightly aft therefrom.
The forward hook 36 includes an arcuate inner lip 36a which is
circumferentially continuous between its opposite ends, and is
spaced radially inwardly from a plurality of circumferentially
spaced apart outer lips 36b spaced radially outwardly from the
inner lip to define a corresponding retention slot 42 therein. The
outer lips 36b are better illustrated in FIG. 3 without the seal 40
illustrated in FIG. 1 supported in the forward hook 36.
The seal 40 illustrated in FIG. 1 includes a sheet metal backing
plate from which the honeycomb seal is supported, with the sheet
metal being bent in the form of a hook which engages the retention
slot 42 in the forward hook 36. The sheet metal hook acts as a
fairing to isolate the seal from purge air in this region in a
conventional manner.
Instead of being circumferentially continuous like the lower lip
36a, the outer lip 36b is segmented for removing substantial
thermal mass from the inner band to reduce or eliminate thermal
mismatch between the forward hook and the inner band during
transient operation. During such operation, the combustion gases 18
flow over the inner band 32 causing heating thereof. Since the
forward hook 36 is hidden below the inner band it is isolated from
the combustion gases and therefore has a delayed thermal response.
By segmenting only the forward hook outer lip 36b, a significant
reduction in thermal mass may be obtained without compromising the
performance of the forward hook for supporting the rotor seal 40 in
a sealed fit therewith.
As shown in FIG. 3, three outer lips 36b are provided, with two
being disposed at the circumferentially opposite ends of the
forward hook, and a single outer lip being disposed intermediate
therebetween, preferably midway. In this way, the fairing portion
of the rotor seal illustrated in FIG. 1 is continuously supported
at its lower surface by the inner lip 36a illustrated in FIG. 3,
and is retained at three points defined by the three outer lips 36b
on its outer surface. Any thermal distortion in the seal fairing
itself during operation is restrained by the three-point retention
on its upper surface and the continuous support at its lower
surface for maintaining effective performance of the fairing. The
reduced thermal mass, however, of the outer lips 36b significantly
reduces thermal mismatch in the inner band with a corresponding
reduction in thermal stress.
Since the individual vanes 28 are integrally joined to both the
outer and inner bands 30,32 as illustrated in FIG. 3, thermal
mismatch therebetween during operation causes locally high thermal
stresses. In accordance with another embodiment of the present
invention, the vanes 28 preferably join the outer band 30 at a
compound radii fillet 44 which extends around the outer perimeter
of the individual vanes.
As shown in more detail in FIG. 4, the fillet 44 preferably has a
larger radius A directly adjacent the radially outer end of the
vane 28 than adjacent the inner surface of the outer band 30. The
large radius A transitions to a smaller radius B in a smooth fillet
blending into the inner surface of the outer band. In an exemplary
embodiment, the compound fillet is effective for significantly
reducing local thermal stress between the vanes and outer band by
about 20%. The radially inner ends of the vanes 28 may join the
outer surfaces of the inner band 32 with conventional single-radius
fillets for acceptable operation, although compound radius fillets
may also be used thereat if desired.
As shown in FIGS. 2 and 3, each of the vanes 28 has an aerodynamic
airfoil configuration including a generally concave, pressure side
28a and a circumferentially opposite, generally convex, suction
side 28b extending axially between leading and trailing edges
28c,d, and radially between the outer and inner bands 30,32 along
the radial span thereof.
As shown in FIGS. 4 and 5, the individual vanes 28 are hollow and
defined by thin walls, and include a radially extending cavity 46
therein. Each vane includes a hollow impingement sleeve or baffle
48 having a plurality of spaced apart impingement holes 50
extending through the thin sheet metal wall thereof. The
impingement holes 50 are preferably arranged in different patterns
facing the inner surface of the vane cavity 46 for preferentially
channeling cooling air 14 in impingement jets toward the midspan of
the vane.
As shown in FIG. 1, the cooling air 14 is bled from the compressor
and suitably channeled through the casing 34 to corresponding inlet
apertures in the top of each vane at the outer band 30 in a
conventional manner. The cooling air 14 initially flows radially
through the inner cavity of the baffle 48 itself, as illustrated in
FIGS. 4 and 5, and then turns for discharge through the impingement
holes 50 against the inner surface of the vane for impingement
cooling thereof.
As shown in FIG. 5, the vane 28 is preferably imperforate over its
entire surface except for a row of trailing edge discharge holes 52
which discharge the spent impingement air along the trailing edge
28d of the vane. The vanes are therefore totally internally cooled
without otherwise providing film cooling holes therethrough or
thermal barrier coating on the external surface of the vane for
avoiding the complexity thereof.
Impingement baffles in turbine nozzles are well known in commercial
use and typically include impingement holes in uniform patterns on
the concave and convex sides thereof. However, in accordance with
another embodiment of the present invention, the pattern of the
impingement holes 50 on the concave side 48a of the baffle as shown
in FIG. 7, corresponding with the pressure side of the vane, is
denser than the pattern on the convex side 48b of the baffle as
shown in FIG. 6, corresponding with the suction side of the vane,
for preferentially cooling the vane midspan.
In the preferred embodiment illustrated in FIGS. 6 and 7, the
impingement holes 50 on both the concave and convex sides 48a,b of
the baffles 48 corresponding with the pressure and suction sides of
the vane preferably have an equal size or common diameter C, which
may be about twenty mils (0.5 mm) for example.
The baffles 48 preferably include imperforate zones or regions 54
at the radially outer and inner opposite span ends thereof that
generally converge toward the baffle midspan between the
corresponding leading and trailing edges 48c,d thereof. In this
way, the improved baffle 48 may use the same amount of cooling air
found in the previous baffle used in commerce, but preferentially
distributes the cooling air to the thermally distressed areas near
the midspan of the vanes.
As shown in FIGS. 5,8 and 11, the baffles 48 preferably also
include a row of larger impingement holes 50b extending along the
leading edges 48c thereof to preferentially cool each vane behind
its leading edge. The large impingement holes 50b illustrated in
FIG. 11 have a diameter D which may be about forty mils (1.0 mm)
for example which is twice the size of the remaining small
impingement holes 50 of diameter C.
The impingement holes 50,50b are preferably arranged in patterns
having different flow density or flow per unit area for
preferentially impingement cooling the different regions of the
vanes. As shown in FIGS. 6 and 9, the impingement holes 50 on the
convex side 48b of the baffle are spaced apart from each other with
a pitch spacing E. Similarly, the impingement holes 50 on the
concave side 48a of the baffle illustrated in FIGS. 7 and 10 have a
pitch spacing F which is preferably smaller than the convex pitch
spacing E for providing increased impingement cooling along the
inner surface of the vane pressure side as opposed to the inner
surface of the vane suction side.
And, in accordance with another embodiment of the present
invention, the large impingement holes 50b along the baffle leading
edge have a greater flow density for preferentially cooling the
inside of the vane leading edge, as shown in FIG. 5, than on the
vane pressure and suction sides. As shown in FIG. 11, the large
holes 50b have a corresponding pitch spacing G which permits the
use of fewer larger holes 50b than smaller holes along the leading
edge for achieving a greater flow density.
It is noted that a given amount of cooling air 14 is provided for
each vane and corresponding baffle which must be suitably
distributed inside the different regions of the vane. The high
density holes on the baffle concave side 48a provide more cooling
of the vane pressure side than the lower density impingement holes
in the baffle convex side 48b on the vane suction side.
Correspondingly, the high density impingement holes 50b along the
baffle leading edge 48c concentrate cooling along the back of the
vane leading edge. The increased amount of impingement cooling air
provided along the vane leading edge and pressure side is at the
expense of a reduced amount on the suction side.
However, by introducing the imperforate regions 54 along both sides
of the baffle near the outer and inner ends thereof, additional
cooling air is provided for the remaining impingement holes by
eliminating impingement cooling in the imperforate regions 54.
As illustrated in FIG. 5, the baffle leading edge 48c has a
relatively large radius of curvature sufficient for forming the
large impingement holes 50b substantially planar therein, and the
vane leading edge 28c has a correspondingly larger radius to
complement the baffle leading edge. In a preferred embodiment, the
radius of the vane leading edge 28c is about 50% greater than in
the previous commercial design for significantly improving the
impingement cooling thereof.
Since the combustion gases 18 stagnate at the vane leading edge
during operation, they effect a correspondingly high external heat
transfer coefficient along the vane leading edge. By increasing the
radius of the leading edge, and correspondingly increasing the
radius of the baffle leading edge 48c and introducing the large
impingement holes 50b therein, a significant increase in the ratio
of the cooling area behind the vane leading edge to the heated area
outside the vane leading edge is provided, with a corresponding
reduction in temperature of the vane leading edge.
FIG. 5 illustrates another embodiment of the present invention
wherein the baffle 48 complements the corresponding vane cavity 46
provided therefor for maintaining a substantially uniform spacing
or gap H with the inner surface of the vane cavity between the vane
leading and trailing edges 28c,d. In this regard, the improved
impingement baffle 48 is longer along its chord or camber axis than
was previously used in the shallow baffle terminating shorter from
both the vane leading and trailing edges.
As shown in FIG. 5, the baffle 48 includes integral standoff pads
56 on its outer surface for spacing the baffle from the opposing
inner surface of the vane cavity 46, with the pads being preferably
arranged on opposite sides of the baffle at both the leading and
trailing edges 48c,d thereof. In this way, both the leading and
trailing edges of the baffle may be maintained in precise alignment
with the corresponding inner surfaces of the vane for maintaining a
substantially uniform gap around the baffle, and in particular at
the vane leading edge which requires enhanced cooling thereof.
As shown in FIGS. 6 and 7, the corresponding pads 56 are more
uniformly spaced on the baffle concave side 48a, corresponding with
the vane pressure side, than on the baffle convex side 48b,
corresponding with the vane suction side. In this way, the vane
pressure side which experiences a relatively high heat transfer
thermal coefficient will enjoy substantially uniform impingement
cooling from the inside by the impingement baffle retained in place
during operation without significant distortion or movement. Since
the vane suction side experiences less heat transfer coefficient on
its external surface, fewer pads 56 may be used on the baffle
convex side with a wider spacing.
These various improvements described above provide tailored and
preferential cooling of the different portions of the nozzle vanes
28 themselves for reducing thermal distress and improving nozzle
durability and life. Furthermore, the improved forward hook 36 of
the inner band 32 and the compound fillet 44 at the outer band 30
provide significant reductions in local thermal stress and mismatch
which further improves the durability and life of the nozzle. The
nozzle therefore enjoys decreased metal temperature during
operation, a more balanced thermal design, and reduced peak
stresses which all directly contribute to increased durability of
the nozzle and enhanced life.
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.
* * * * *