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Publication numberUS3730640 A
Publication typeGrant
Publication dateMay 1, 1973
Filing dateJun 28, 1971
Priority dateJun 28, 1971
Publication numberUS 3730640 A, US 3730640A, US-A-3730640, US3730640 A, US3730640A
InventorsIverson V, Pedersen H, Rice A
Original AssigneeUnited Aircraft Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Seal ring for gas turbine
US 3730640 A
Abstract
A gas turbine blade tip ring seal, which surrounds the blade tips to minimize turbine gas leakage around the blades, that incorporates a thermal protection system that allows modulation of the seal leading edge temperature level and thermal response to minimize seal deterioration during turbine transient operating conditions.
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Description  (OCR text may contain errors)

I United States Patent 1 1 3,739,646

Rice et a1. May 1, 1973 [54] SEAL RING FOR GAS TURBINE 3,367,628 2/1968 Fittonm, ..415/1 15 2,984,454 5/1961 Fiori ..415/117 [75] Inventors: Alvin S. Rice, South Glastonbury;

Herbert C. Pedersen, Vernon, both FOREIGN PATENTS OR APPLICATIONS of Conn.; Vincent L. Iver-son, Palm Beach Gardens Fla 601,410 1/1960 Italy ..4l5/117 [73] Assignee: United Aircraft Corporation, East Primary ExaminerHenry F. Raduazo Hartford, Conn. Att0rney-Charles A. Warren [22] Filed: June 28, 1971 ABSTRACT [21] Appl' 157515 A gas turbine blade tip ring seal, which surrounds the blade tips to minimize turbine gas leakage around the [52] U.S. Cl. ..415/117, 415/172, 415/217 blades, that incorporates a thermal protection system [51] Int. Cl ..'..F01d 11/08 that allows modulation of the seal leading edge tem- [58] Field of Search ..415/115, 116, 117, perature level and thermal response to minimize seal 415/175, 176, 171, 172, 217; 60/3966 deterioration during turbine transient operating conditions.

56 R ferences C't d 1 e l e 6 Claims, 2 Drawing Figures UNITED STATES PATENTS 3,451,215 6/1969 Barr ..4l5/1l6 SEAL RING FOR GAS TURBINE BACKGROUND OF THE INVENTION Because of temperature differentials in the gas turbine, a seal that surrounds a row of blades in the turbine particularly the first stage row of blades is subjected to significant temperature changes during engine operation and particularly during transient conditions. Under steady state conditions, this seal which regularly extends between the shrouds on adjacent rows of vanes and thus isexposed directly to the hot. turbine gas between the shrouds of one row of vanes and the adjacent shrouds in the blade has a temperature difference between upstream and downstream edges of the seal ring of as much as 400F. Under transient conditions this temperature difference becomes much greater to such an extent that the seal ring is subject to plastic deformation and when it cools, it shrinks to a smaller dimension than originally with resultant reduction in the clearance around the blade tips. After several operating cycles the shrinkage is enough to cause damaging rubbing of the seal and blade tip shrouds.

SUMMARY OF THE INVENTION One feature of the invention is an arrangement for shielding the seal ring to minimize the amount of heat absorbed by radiation from the hot turbine inlet guide vanes and convection from the turbine gas path and also to guide cooling air over the surface of the ring thereby keeping the ring cooler particularly at the upstream portion which is exposed to the highest turbine gas temperature. Another feature is the incorporation of this shield in the turbine structure with a minimum of engine modification and at the same time obtain the desired flow of cooling air into the space between the seal ring and the overlying shield.

In accordance with the invention, a shield is positioned in the. space between the downstream edges of the shrouds in the turbine vanes and the tips of the adjacent turbine blades such that this shield overlies the seal ring in this area in spaced relation thereto. This shield may also incorporate an integral ring surrounding the vane shrouds for assisting in directing adequate cooling air into the space between the seal ring and the overlying shield thereby guiding coolingair over the seal ring at its hottest portions.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a fragmentary longitudinal sectional view through the first stage of the gas turbine.

FIG. 2 is a fragmentary plan view of a portion of the shield.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring first to FIG. 1 the invention is shown in connection with the first turbine stage of a multi-stage gas turbine, one example of which is shown in the Savin US. Pat., No. 2,747,367. As shown, the turbine has a casing 2 having an inwardly extending flange 4 which supports the outer shrouds 6 of the row of inlet guide vanes 8 for the turbine. These vanes are hollow and air 1 cooled as will be described. Each shroud has an outwardly extending flange 10 secured as by bolts 12 to the casing flange 4. A sleeve 14 surrounding the bolts 12 serves to determine the spacing of the nuts 16 from the casing flange 4 for a purpose which will appear.

The downstream edges of the shrouds abut against a Z-shaped ring 18 which in turn rests against a second flange 20 on the casing 2, this second flange being spaced downstream from the first flange. This Z ring extends forwardly in surrounding relation to the row of turbine vanes outwardly of the shrouds 6 and has an outwardly extending flange 22 adjacent the flange 4 and positioned between adjacent bolts 12. The second flange 20 has a plurality of passages 24 therethrough for cooling air.

Downstream of the row of vanes are the cooperating rotating blades 26 carried on a disc 28. These blades have tip shrouds 30 thereon which may have outwardly projecting fins 32. Cooperating with these shrouds and fins is the surrounding seal ring 34 which extends from a point adjacent to flange 20 downstream beyond the blade shrouds 30 to overlie a third flange 36 on the casing. The seal 34 ring and flange 36 have cooperating loose splines 38 and 40 for locating the seal ring in surrounding relation to the blades but with a substantial freedom of radial movement to provide for expansion. The seal ring is spaced from flange 20 at its upstream end to provide a clearance and is surrounded by a spring 42 the configuration of which serves to hold the ring in a position to maintain this clearance. This spring also serves as a vibration damper for the seal ring.

The axial dimension of the seal ring is such that it extends beyond the blade shrouds in both upstream and downstream directions so that the ring would regularly be exposed directly to the hot gas passing through the turbine. Under these circumstances the seal ring would be exposed upstream of the blade shrouds to the temperature of the gas between the vane shrouds and the blade shrouds, and this gas under steady state conditions may be 400 hotter than the gas temperature downstream of the blade shrouds to which the downstream portion of the seal ring is exposed. .Under transient conditions this temperature difference will be significantly greater. These temperature differentials subject the ring to plastic deformation which results in dimensional changes such as shrinkage of the ring with successive operating cycles of the engine.

To minimize these thermal changes, the Z ring 1 which has an inwardly extending flange 44 engaging against the upstream face of flange 20 also has a cylindrical flange or shield 46 extending downstream over the end of flange 20 and into overlying relationship with the upstream portion of seal ring 34. This shield 46 extends a substantial part of the distance toward the row of blade shrouds 30 preferably almost to these blade shrouds. This should serve to shield the underlying portion of the seal ring from direct contact with the hot gases in this area and from radiation from the hot vanes. Cooling air through the passages 24 flows around the upstream end of the seal ring and is guided by the shield over the inner surface of the seal ring. The result is to reduce the temperature differentials in the seal ring to such an amount as to reduce the seal ring shrinkage to about one-tenth of the previous shrinkage.

The cooling air used is obtained from around the flame tubes and is referred to as secondary combustion air. This air is in the space 48 upstream of casing flange 4 and passes between flange 22 and flange 4 into the space 50 between casing flanges 4 and and outside of Z ring 18. This cooling air then flows through passages 24 into space .52 downstream of flange 20 and around the upstream end of the seal ring 34 as above described. Some of the cooling air in space 48 also enters the space 54 radially inward of the Z ring 18 and enters the vanes 8 for cooling them. Some of this cooling air also flows inwardly between adjacent shrouds 6 to flow over and cool the inner surface of these shrouds and also over the inner surface of the shield for cooling this to some extent.

The Z ring 18 is preferably slotted to minimize thermal distortion. As shown in FIG. 2, deep slots 56 extend forwardly from the downstream edge of shield 46 to include flange 44 and into the main portion of the Z ring. The forward end of the Z'ring is also preferably notched as at 58 and these notches also extend into the main portion of the Z ring. These notches are relatively wide and provide passages for the flow of cooling air from space 48 into the space 54 as above described.

We claim:

1. The combination with a row of turbine vanes having outer shrouds, a casing surrounding the outer shrouds and supporting said row of vanes, a row of turbine blades downstream of and adjacent to said varies and a continuous seal ring extending around the tips of the blades in close proximity thereto and supported within said casing, said seal ring also projecting forwardly of the tips of the blades, of a shield extending downstream of and separate from the shrouds and substantially in alignment therewith and overlying at least a part of the seal ring for shielding said ring from the hot gases passing over the van'es and blades.

2. The combination as in claim 1 in whichthe shield extends substantially to the blade tips;

3. The combination as in claim 1 in which means are provided for discharging cooling air into the space between the seal ring and the shield.

4. The combination as in claim 1 in which the seal ring extends upstream substantially to the shrouds and the shield overlies this portion of ring and extends substantially to the tips of the blades. 4

5. The combination with a row of turbine vanes having outer shrouds, a row of turbine blades downstream of and adjacent to said varies, a casing surrounding said vanes and blades and supporting said vanes, and a continuous seal ring extending around the tips of the blades in close proximity thereto and also projecting forwardly of the tips of the blades substantially to the vane shrouds, said seal ring being supported by the casing for radial movement relative thereto, of a shield extending downstream of and separate from the shrouds and substantially in alignment therewith and overlying the portion of the seal ring between the vaneshrouds and blade tips for shielding said ring.

6. The combination as in claim 5 in which the casing has an inwardly extending flange for supporting the shield, said flange having means therein for directing cooling fluid to the outer surface of the upstream end of the seal ring.

' a: is

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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3885886 *Jun 27, 1973May 27, 1975Mtu Muenchen GmbhUnshrouded internally cooled turbine blades
US3893782 *Mar 20, 1974Jul 8, 1975Westinghouse Electric CorpTurbine blade damping
US3992126 *Mar 25, 1975Nov 16, 1976United Technologies CorporationTurbine cooling
US4053254 *Mar 26, 1976Oct 11, 1977United Technologies CorporationTurbine case cooling system
US4190397 *Nov 23, 1977Feb 26, 1980General Electric CompanyWindage shield
US4242042 *May 16, 1978Dec 30, 1980United Technologies CorporationTemperature control of engine case for clearance control
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Classifications
U.S. Classification415/117, 415/173.6, 415/180
International ClassificationF01D11/08
Cooperative ClassificationF01D11/08
European ClassificationF01D11/08