U.S. patent number 3,864,056 [Application Number 05/383,426] was granted by the patent office on 1975-02-04 for cooled turbine blade ring assembly.
This patent grant is currently assigned to Westinghouse Electric Corporation. Invention is credited to Frank K. Gabriel, Stephen D. Leshnoff.
United States Patent |
3,864,056 |
Gabriel , et al. |
February 4, 1975 |
COOLED TURBINE BLADE RING ASSEMBLY
Abstract
A cooled turbine blade ring assembly surrounds an alternating
arrangement of stator blades and rotating blades. A cooling air
chamber directs cooling fluid to each stator vane retainer
arrangement and to each ring segment which is circumferentially
disposed about each rotor disc. The ring segments have a heat
shield disposed on their outward side. The stator vane retainer
arrangement including the ring segments and heat shields are
maintained in a controlled temperature and in a controlled
pressurized state. A bias producing means holds the heat shield and
ring segments in a radially inwardly directed position, against the
blade ring, to which it transfers reaction loads. The flow of
cooling fluid maintains the heat shield and ring segments at a
controlled temperature, while it prevents overcooling of the outer
shroud portions of each vane. The heat shield also protects the
blade ring and associated cooling air mechanism, and prevents
damage to this structure due to distortion by thermal gradients
within certain portions of the ring segments and blade ring.
Inventors: |
Gabriel; Frank K. (Springfield,
PA), Leshnoff; Stephen D. (Highland Park, NJ) |
Assignee: |
Westinghouse Electric
Corporation (Pittsburgh, PA)
|
Family
ID: |
23513094 |
Appl.
No.: |
05/383,426 |
Filed: |
July 27, 1973 |
Current U.S.
Class: |
415/178; 415/177;
415/116; 415/136 |
Current CPC
Class: |
F01D
9/041 (20130101); F01D 5/08 (20130101); F01D
11/08 (20130101); Y02T 50/60 (20130101) |
Current International
Class: |
F01D
11/08 (20060101); F01D 5/02 (20060101); F01D
9/04 (20060101); F01D 5/08 (20060101); F01d
025/08 (); F01d 025/14 (); F01d 025/12 () |
Field of
Search: |
;415/115,116,117,136,137,178,177 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
|
|
|
|
|
|
|
1,219,504 |
|
May 1960 |
|
FR |
|
1,020,900 |
|
Feb 1966 |
|
GB |
|
Primary Examiner: Freeh; William L.
Assistant Examiner: Casaregola; L. J.
Attorney, Agent or Firm: Telfer; G. H.
Claims
We claim:
1. A hot elastic fluid machine, comprising: a turbine casing, a
plurality of rotatable discs mounted on an axis, a plurality of
rotor blades disposed on the periphery of each of said rotatable
discs, a plurality of radially directed stationary blades disposed
in annular arrays alternating with said rotatable discs, each
annular array of stationary blades having an inner and an outer
shroud ring, a blade ring circumferentially disposed about the
blades, means for supplying said stationary blades with cooling
fluid, a plurality of insulated ring segments coaxial with said
blade ring disposed radially outwardly of said rotating blades and
inwardly of said blade ring, conduit means for supplying
pressurized cooling fluid to said ring segments, said ring segments
disposed about said rotating blades being expansible in both the
radial and the axial direction, said radial direction of expansion
of said ring segments being controlled by a biasing means, and a
heat shield, said heat shield being disposed in a spaced
relationship and radially outwardly of, yet fixedly attached to,
said ring segments, said heat shield being cooled by chambers
having a passageway extending therethrough, said passageway
permitting the flow of cooling fluid through said bias producing
means, the radially innermost chamber having holes therein, said
holes permitting the pressurized cooling fluid to impinge upon said
heat shield, the flow of cooling fluid upon said heat shield
preventing localized overcooling and hence preventing localized
distortion of said ring segments thereby, said heat shield
receiving heat from said ring segments by radiation and convection
to permit a uniform temperature across said ring segments and, said
heat shield providing a heat barrier between said blade ring and
said ring segments.
2. A hot elastic fluid turbine machine as recited in claim 1,
wherein said bias producing means are generally radially directed
springs, wherein said cooling fluid passes through said springs and
impinges upon said heat shield adjacent said ring segments, said
heat shield preventing heat from penetrating into said blade ring
area which is comprised of less heat resistant material.
3. A hot elastic fluid turbine machine as recited in claim 1,
wherein said expansible ring segments are provided with both
upstream and downstream support means, said support means consists
of a plurality of circumferentially disposed vane segments, each of
said vane segments having circumferentially directed shoulder
portions, said shoulder portions overlapping lip portions of said
ring segments, said lip and shoulder portions being in sliding
contact, and at the interface between said lip and shoulder
portions there is a plurality of passageways disposed thereacross,
said passageways permitting cooling of said lip and shoulder
portions, and said passageways providing escape means for the spent
cooling fluid.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to gas turbines, and more
particularly, to cooled blade ring assemblies disposed about the
blades of the gas turbine.
2. Description of the Prior Art
In gas turbines, designers try to avoid the undersirable effects of
thermal expansion and heat upon the blade ring and the housing
disposed about the stationary and the rotating blades. This is
generally done by supplying cooling air from a compressor and
directing it through the stationary vanes. This air is also usually
directed radially outwardly from an axially disposed source through
a rotating rotor disc and out through the rotating blades which are
also cooled by this air. The cooling air passing out the rotating
blades usually cools the ring segments disposed radially outwardly
thereof, but this is undesirable in that it does not have the
necessary cooling effect to prevent thermal distortion and buckling
of members therein. The rotating blades themselves at the
downstream end of the turbine are not cooled, leaving the blade
rings at that end, with little or no cooling.
Other attempts have been made to direct the cooling fluid radially
inwardly upon the blade ring structure from the compressor source,
but this does not protect the shielding from the hot gaseous
fluid.
Some of the prior art devices merely cause a cooling fluid to flow
over the radially outer surface of the blade ring. This cooling
technique does not provide temperature regulation nor does it
provide adequate protection for the blade ring and surrounding
stationary turbine containment from the hot gases flowing in the
hot motive fluid flow path.
SUMMARY OF THE INVENTION
In accordance with one embodiment of the invention, a plurality of
blade ring segments and vane shroudings are disposed
circumferentially about the turbine axis and rotor. The blade ring
segments and stationary flow path segments define an annular
cooling fluid flow chamber, from which cooling fluid is channeled
through passageways into each stator vane assembly. Also, cooling
fluid is passed under pressure through passageways from the annular
chamber to impinge upon a heat radiation shield which is attached
to each of a plurality of circumferential disposed ring segments.
The heat shield cooperates with a spring member that maintains the
radiation shield and its attached ring in a generally stationary
yet expendable position. The cooling fluid passes around the shield
creating a pressure differential between the cooling air chamber
and the hot motive fluid gaseous flow path, preventing the hot
working gaseous fluid from entering and damaging the cooling air
chamber and its contiguous blade ring structure.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a diagrammatic view of a gas turbine engine embodying
this invention, with parts broken away;
FIG. 2 is an enlarged view of the cooled blade ring assembly of
FIG. 1;
FIG. 3 is one embodiment of a ring segment cooling arrangement;
and
FIG. 4 is a perspective view of the radiation shield.
DESCRIPTION OF THE PREFERRED EMBODIMENTS
Referring to the drawings, particularly to FIG. 1, the structure
shown therein comprises an axial flow gas turbine 10 having three
rotor discs 12, 14 and 16, each having mounted on their respective
peripheries a plurality of rotating blades 18, 20 and 22. Each of
the rotating blades 18, 20 and 22 has an air foil portion 23. A
nozzle or row one array of stationary vanes 24 is disposed upstream
of the first array of the rotating blades 18. Between the first two
arrays of rotating blades 18 and 20 there is disposed another
circumferential array of stationary blades 26. Downstream of the
second array of rotating blades 20 there is disposed another array
of circumferentially disposed stationary vanes 28, followed by
another array of rotating blades 22. A combuster 30 supplies the
hot motive fluid which passes through the nozzle 24 and into the
blades 18, 20 and 22, thereby causing their respective rotors 12,
14 and 16 to rotate.
The hot motive fluid travels through an annular flow path, defined
on its radially inward portion by a plurality of shroud rings 32
supporting the radially inner portion of the stationary vanes 24,
26 and 28 and by a plurality of blade platforms 34. A blade
platform 34 is provided on each radially inward portion of each air
foil 23. The hot motive fluid flow path is defined on its radially
outward portion by a plurality of circumferentially disposed outer
vane segments 35, a blade ring 36, and a plurality of ring segments
37. The entire turbine 10 is enclosed in a turbine cylinder 38,
only a portion of which is shown in FIG. 1.
High efficiency and high power output of the turbine requires
cooling of the surfaces of the boundaries defining the hot motive
fluid flow path. Without cooling of these surfaces, the entire
turbine structure would have to be constructed of a more exotic,
more expensive material, if such materials are even available, that
would withstand the high thermal effects of the more powerful
turbine.
Cooling is performed in this invention, by providing a cooling
inlet 40 disposed through the turbine cylinder 38. Pressurized
cooling fluid is supplied to the inlet 40 from a source, not shown.
This cooling fluid is received in an annular chamber 42 defined by
the blade ring 36. The annular chamber 42, for receiving the
cooling fluid, is shown more clearly in FIG. 2.
A ring segment cooling arrangement 44 is shown more clearly in
FIGS. 2 and 3. The ring segment cooling arrangement 44 is shown
disposed circumferentially about the hot motive fluid flow path,
radially outwardly of the rotating arrays of the blades 18 and
20.
The ring segment cooling arrangement 44 includes the ring segments
37, which are positioned in an annular array around and adjacent
the hot motive fluid flow path. A heat shield 46, shown in FIG. 4,
is attached to the radially outer side of each ring segment 37. The
shield 46 comprises a corrugated member 48 welded to the radially
inward side of a sheet member 50. Only a central linear segment of
the shield 46 is welded to the ring segment 37. This allows for
expansion due to thermal effects for the individual portions of the
heat shield 46. The corrugated member 48 prevents much of the heat
transfer from the hot fluid flow path to the blade ring 36, and
therefore, acts as an insulator.
The heat shield 46 is cooled by an impingement of cooling fluid
ejected from a plurality of holes or jets 51 disposed in the
radially inner end of a radially directed chamber 52. The chamber
52 acts as a cooling fluid conduit. The chamber 52 is held in a
compressive state against the shield 46 by a spring 54. The spring
54 and chamber 52 cooperate with and are disposed within a
sleeve-like tube 56 that is attached to the blade ring 36.
The radially outer end of tube 56 is disposed within the cooling
fluid chamber 42 of the blade ring 36. The cooling fluid is forced
in a cooling fluid inlet orifice 57 on the outer end of tube 56.
The cooling fluid passes through the spring member 54, and into the
chamber 52. The fluid is then forced out the openings 51 in the
radially inner end of chamber 52. The fluid impinges upon and flows
over the shield 46. The spent cooling fluid escapes around the
axially disposed portion of the ring segments 37 through a
plurality of gaps 58 which are disposed circumferentially around
the ring segments 37 between the ring segments 37 and the
supporting adjacent vane segments 35, and/or the fluid escapes
through an annular array of passageways 65 disposed in shoulder
portion 35A, as shown in FIG. 3.
As the cooling fluid passes into the hot motive fluid flow path, it
creates a pressure differential between the motive fluid flow path
and the coolant fluid flow passageways. This differential in
pressure prevents any hot motive fluid from entering the ring
segment cooling arrangement 44 or the blade ring structure 36.
The spring member 54 holding the chamber 52 and the shield 46 in a
compressed state of engagement with the vane segments 35, allows
for thermal growth in a radial direction of each of the ring
segments 37 and their respective shield members 46. Axial changes
due to thermal expansion are permitted because of an overlap
arrangement between the axially disposed, circumferentially
directed lip portions 37A of the ring segments 37 and the axially
disposed circumferentially directed shoulder portions 35A of the
adjacent vane supporting segments 35.
The vane supporting segments 35 themselves are cooled by the flow
of pressurized cooling fluid entering a plurality of
circumferentially disposed orifices 59 in the blade ring 36 as
shown in FIG. 2. The orifices 59 supply the cooling fluid from the
cooling fluid chamber 42 within the blade ring 36 and permit the
cooling fluid to pass through the first array of stationary vanes
24.
The coolant fluid entering the annular arrays of downstream
orifices 59' provide cooling for the stationary vanes 26 and 28, as
shown in FIG. 2. The fluid impinges upon a splash shield 60 that
directs the cooling fluid to flow near the axially disposed
portions of each of the vane segments 35. The axially disposed
portions of the stationary vane segments 35 are near those members
that provide support for the ring segments 37.
This cooling effect minimizes thermal growth of the vane segments,
the blade rings and the supporting structure, and permits the
manufacture of a high performance turbine with a satisfactory metal
creep life from a more common, less heat resistant and less
expensive metal.
From the foregoing description, it is apparent that many
modifications may be made by those skilled in the art. For
instance, the ring segments could be compressed by a different
arrangement of bias producing members. The cooling fluid could be
ejected from several tubes extending from the cooling air chamber;
and the heat shield is susceptible to various modifications.
* * * * *