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Publication numberUS3864056 A
Publication typeGrant
Publication dateFeb 4, 1975
Filing dateJul 27, 1973
Priority dateJul 27, 1973
Also published asCA995140A, CA995140A1, DE2432092A1
Publication numberUS 3864056 A, US 3864056A, US-A-3864056, US3864056 A, US3864056A
InventorsGabriel Frank K, Leshnoff Stephen D
Original AssigneeWestinghouse Electric Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Cooled turbine blade ring assembly
US 3864056 A
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.
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Description  (OCR text may contain errors)

United States Patent 1191 Gabriel et al.

Feb. 4, 1975 COOLED TURBINE BLADE RING Primary Examiner-William L. Freeh ASSEMBLY Assistant Examiner-L. J. Casaregola [75] Inventors: Frank K. Gabriel, Springfield, Pa.; Attorney Agent or Flrm G Telfer SNteIphen D. Leshnoff, l-l1ghland Park, ABSTRACT 1 A cooled turbine blade ring assembly surrounds an al- [73] Asslgnee' Westmghuuse Elecmc Corporal, ternating arrangement of stator blades and rotating Pmsburghi blades. A cooling air chamber directs cooling fluid to [22] Fil d; J l 27 1973 each stator vane retainer arrangement and to each ring segment which is circumferentially disposed [21] Appl' 383,426 about each rotor disc. The ring segments have a heat shield disposed on their outward side. The stator vane [52] us. c1 415/178, 415/177, 415/116 retainer arrangement including the ring Segments and [51] I Cl Fo 25/0 o 25/14 o 25/12 heat shields are maintained in a controlled tempera- [5 Fi ld f Search 415 15 116, 7 3 ture and in a controlled pressurized state. A bias pro- 415 137 17 177 ducing means holds the heat shield and ring segments in a radially inwardly directed position, against the [56] References Cited blade ring, to which it transfers reaction loads. The UNITED STATES PATENTS flow of cooling fluid maintains the heat shield and ring 2427 244 9 9 w segments at a controlled temperature, while it pre- 3034298 x 2; ig? vents overcooling of the outer shroud portions of each 3,236,069 5/1973 Beam et a] 415/173 vane- 3,298,823 1/1967 Waugh 415/136 The heat shield also protects the blade ring and 3,451,215 6/1969 Barr 415/116 associated cooling air mechanism, and prevents FOREIGN PATENTS OR APPLICATIONS damage to this structure due to distortion by thermal 1 219 504 5/1960 France 415/136 gradients within certain portions of the ring segments 1,020,900 2/1966 Great Britain 415/136 and blade 3 Claims, 4 Drawing Figures .57 Ll 2 Q a Q, \3

5 Q 4 P .4 51 51 3 46 J 35 y i;

Kat/11m 1 65 35 35A R58 3? 37A 35A 58 1 COOLED TURBINE BLADE RING ASSEMBLY BACKGROUND OF THE INVENTION 1. Field of the Invention DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring to the drawings, particularly to FIG. I, the

This invention relates generally to gas urb es. an structure shown therein comprises an axial flow gas turmore 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 undesirable 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 vaneassembly. 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.

bine 10 having three rotor discs l2, l4 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 I8. 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 withstandthe 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 52v cooperate with and are disposed within a sleeve-like tube 56 that is attached to the blade ring 36.

Theradially 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.

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.

Patent Citations
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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3992127 *Mar 28, 1975Nov 16, 1976Westinghouse Electric CorporationStator vane assembly for gas turbines
US4019320 *Dec 5, 1975Apr 26, 1977United Technologies CorporationExternal gas turbine engine cooling for clearance control
US4087199 *Nov 22, 1976May 2, 1978General Electric CompanyCeramic turbine shroud assembly
US4317646 *Feb 22, 1980Mar 2, 1982Rolls-Royce LimitedGas turbine engines
US4379677 *Oct 7, 1980Apr 12, 1983Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A."Device for adjusting the clearance between moving turbine blades and the turbine ring
US4462204 *Jul 23, 1982Jul 31, 1984General Electric CompanyGas turbine engine cooling airflow modulator
US4551064 *May 24, 1985Nov 5, 1985Rolls-Royce LimitedTurbine shroud and turbine shroud assembly
US4635332 *Sep 13, 1985Jan 13, 1987Solar Turbines IncorporatedSealed telescopic joint and method of assembly
US5116199 *Dec 20, 1990May 26, 1992General Electric CompanyBlade tip clearance control apparatus using shroud segment annular support ring thermal expansion
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US7665958Sep 24, 2007Feb 23, 2010Alstom Technology Ltd.Heat accumulation segment
US9335051 *Jul 13, 2011May 10, 2016United Technologies CorporationCeramic matrix composite combustor vane ring assembly
US9534607 *Mar 2, 2011Jan 3, 2017Kawasaki Jukogyo Kabushiki KaishaCompressor of use in gas turbine engine
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Classifications
U.S. Classification415/178, 415/177, 415/136, 415/116
International ClassificationF01D5/02, F01D5/08, F01D11/08, F01D11/10, F01D9/04
Cooperative ClassificationF01D9/041, F01D5/08, F01D11/08
European ClassificationF01D5/08, F01D11/08, F01D9/04B