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Publication numberUS3742705 A
Publication typeGrant
Publication dateJul 3, 1973
Filing dateDec 28, 1970
Priority dateDec 28, 1970
Publication numberUS 3742705 A, US 3742705A, US-A-3742705, US3742705 A, US3742705A
InventorsP Sifford
Original AssigneeUnited Aircraft Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Thermal response shroud for rotating body
US 3742705 A
Abstract
This invention shows a shroud construction located around the tips of the blades on a rotating body in an engine to provide a minimum clearance between the blade tips and the shroud during all conditions of operation-acceleration, steady state and deceleration. This shroud construction is connected to the end of two annular flanges which extend inwardly from an inner engine case towards the rotating member. The shroud member faces the blade tips and is arranged to have a flow of cooling fluid pass therethrough. Cooling vanes extend from the inner engine case into a duct through which fan air passes. Compressor air passes through the shroud construction.
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Description  (OCR text may contain errors)

United States Patent 11 1 Sifiord [75] lnventorz Perry P. Sifford, Palm Beach Garden, Fla.

[73] Assignee: United Aircraft Corporation, East Hartford, Conn.

[22] Filed: Dec. 28, 1970 v [21] Appl. No.: 101,481

[52] US. Cl. 60/39.66, 415/117 [51] Int. Cl. F02g 1/00 [58] Field of Search 60/3966, 262, 226, 60/266; 415/110, 115, 116, 175, 144, 196, 117

[56] References Cited UNITED STATES PATENTS 2,880,574 -4/1959 Howald 60/262 2,859,934 ll/l958 Halford..... 415/115 3,391,904 -7/l968 Albert 415/115 THERMAL RESPONSE SHROUD FOR ROTATING BODY 1451 July 3, 1973 3,451,215 6/1969 Barr 415/116,

Primary Examiner-Samuel Feinberg Attorney-Jack N. McCarthy 1. I I ABSTRACT This invention shows a shroud construction located around the tips of the blades on a rotating body in an 5 Claims, 2 Drawing Figures PAIENIEnJuLa ms 3.742.705

THERMAL RESPONSE SHROUD FOR BODY BACKGROUND OF THE INVENTION This invention relates to a device for minimizing the clearance between blade tips and surrounding seal. In this art, many different types of shrouds have been used. A sample of these are shown by US. Pat. No. 3,391,904; 2,859,934; and 3,443,791.

SUMMARY OF THE INVENTION A primary object of this inveniton is to provide a simplifled thermal response shroud for location around turbine blade tips to obtain good performance. This invention provides a close steady state tip clearance without detrimental rubbing during engine transients.

In accordance with the present invention, clearance at room temperature is large but the steady state clearance is kept low by increased cooling in the case and shroud support structure.

In accordance with a further aspect of the present invention, the shroud provides adequate tip clearance during deceleration by retarding the cooling rate of the shroud support structure.

This invention provides a shroud support having a mass which is greater than cooling fins which are arranged to remove heat therefrom. The shroud is formed of two members between which fluid passes to cool the shroud. Cooling fins extend outwardly from ROTATING shroud support members to another source of cooling fluid.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a representative showing of an aircraft gas turbine fan engine with a section broken away showing the location of the invention.

FIG. 2 is an enlarged view of the area shown by the broken away section in FIG. 1.

DESCRIPTION OF THE PREFERRED EMBODIMENT Referring to FIG. I a gas turbine power plant 1 is shown of the fan type. The power plant has a fanand compressor section 2, a combustion section 4, a turbine section 6 and an exhaust section 8. The turbine section 6 is shown broken away to locate the invention. An enlarged view of this is shown in FIG. 2 andincludes an outer casing 10 and an inner casing 12. The inner casing 12 adjacent the outer tips of the turbine blades 14 is formed having a forward inwardly extending annular flange 16 and a rearward inwardly extending annular flange 18 which supports a shroud construction 20 which encompasses the ends of the tips of the blades 14.

Stator vanes 22 have their rear ends outside of their outer platforms abutting the forward surface of the annular flange 16 at 26. Stator vanes 24 have their front ends outisde of their outer platforms connected to and spaced from the rearward surface of the annular flange I8 in a manner to be hereinafter described. Specific connections of vanes in gas turbine engines are shown in U. S. Pat. Nos. 3,295,824; 3,391,904; 3,423,071 and Annular flange 16 has two rearwardly projecting cylindrical-like flanges 40 and.42. Flange 40 extends from the end of the flange 16. A cylindrical-like groove 44 is formed between said flanges. The inneredge of the outer cylindrical flange 42 is undercut at 46 so that a cylindricaljinner surface 48 appears opposite and at the same radius as the inner surface 50 of the top of' flange l8. Annular flange 18 has two rearwardly pro-' jecting cylindrical-like flanges and 62 having a groove 64 formed therebetween. Flange 60 extends from the end of the flange 18. A cylindrical groove 34 is formed between the flange 62 and a portion of the inner casing l2. Stator vanes 24 have a flange 28 projecting inwardly therefrom with a plurality of notches 30 cut therein around the entire circumference. Flanges 32 extend forwardly from the outer end of the flanges 28 to engage the groove 34 of the flange 18.

A cylindrical liner 52 is pressed into position contacting cylindrical surfaces 48 and 50 within inner case 12 for a purpose to be hereinafter described. Liner 52 has a plurality of holes 51 located around its forward end adjacent the rear edge of flange 42.-A rub strip seal member 54 is formed as a cylindrical-like member made in segments having an inner seal face 56. Segmenting permits rub strip thermal growth without changing the ring diameter.

The forward end of said rub strip member 54 has a forwardly extending portion 61 of reduced diameter which projects into the opening 44. The rearward end of said rub strip member 54 has an outwardly extending flange 66 which has an annular groove 68 therein on its forward face. The annular groove 68 forms a forwardly projecting cylindrical flange 70 at the outerend of the flange 66. Rub strip member 54 is positioned over the liner 52 with the forward portion 61 projecting into the opening 44 and with the flange 60 extending into the groove 68 and flange 70 extending into the groove 64.

The inner face of the rub strip member 54 is formed having grooves 63 and fins thereby forming a plurality of passages with the liner 52 whenit is in place. An annular groove 90 is cut in the strip member 54 and connects the forward end of all of the grooves 63. An annular groove 92 is cut atthe rearward end to connect all of the grooves 63. Holes 51 permit the flow of fluid from chamber 84 into the manifold formed by the groove 90 and liner 52. Rub strip member 54 has a plurality of passageways 94 located therethrough at its rearward end which connects the groove 92 to the exterior of said rub strip member at a point just forwardly of the outer platforms of the vanes 24.

A plurality of holes extend through the flange 16 between the inner casing 12 and the outer ends of the vanes 22. These openings admit compressor air from the chamber 82 to pass into the annular chamber 84 formed by the inner casing 12 and its two annular flanges I6 and 18 along with the liner 52. Openings 86 are located in the flange 18 in a manner to permit flow from the chamber 84 through the notches 30.

Holes 80, openings 86, and holes 51 are sized to obtain a desired flow through the grooves 63 forming pas- 'sages with the liner 52 for discharge onto the outer ends of the vanes 24. Cooling fins 96 and 98 extend around the circumference of the casing l2 and project into the outer duct carrying fan air past the turbine section.

Cooling flow regulation depends on engine operating conditions and is a function of the fan and compressor speed. When the engine accelerates, fan and compressor cooling air flow increases and when the engine decelerates, the fan and compressor cooling air flow decreases. During steady state operation, fins 96 and 98 are heated by case 12 and flanges l6 and 18, and cooled by fan cooling air, resultant temperature being relatively low. Likewise, flanges 16 and 18 are heated throughconduction from rub strip 54 and cooled by compressor air flowing through holes 80 and 86 with resultant temperature being relatively high.

During deceleration, cooling capacity decreases as described above. Heating rate from rub strip 54 also decreases. Reduction cooling flow allows residual heat in flanges 16 and 18 to transfer, by conduction, to fins 96 and 98, the temperature of said fins thereby increasing, and overall cooling rate of the structure decreasing. The heat transfer rate from the annular flanges 16 and 18, the shroud support, depends on its mass and heat capacity which is greater than that of the cooling flns 96 and 98. The greater the mass of the flanges 16 and 18, the greater the residual heat and thus the slower the overall cooling rate. The result of the reduced overall cooling rate during deceleration is reduction of the rate of radially inward movement of the shroud. This arrangement provides a shroud construction 20 movement that responds to engine operating conditions to maintain a minimum blade tip and shroud clearance gap to avoid deceleration rubbing.

During steady state operation, for a fan engine tested, the gas stream temperature flowing over the blade tips was 2,250F while the compressor cooling air passing through the openings 80 was approximately l,020F and the temperature of the air passing the fins 96 was approximately 600F.

I claim:

1. In an engine, a compressor, combustion means producing hot gases, a turbine motor having blades extending about the periphery thereof, said blades being arranged so that hot gases from said combustion means flow thereover during operation, shroud means extending around the outer tips of said blades for providing a seal therewith, said shroud means receiving heat from the hot gases of said combustion means, an inner engine casing encompassing said shroud means, said inner 4 casing having inwardly extending flange means, said shorud means being connected to said flange means for radial movement therewith during thermal changes, an outer casing being spaced from and extending around said inner casing, cooling fin means extending outwardly from said inner casing for controlling the temperature of the inwardly extending flange means,

means for discharging a flow of cooling air over said cooling fin means to maintain a close blade tip clearance with said shroud meanss during steady state operation of said engine, said cooling fin means having a mass which is less than the flange means to reduce the overall cooling rate of the flange means when said engine decelerates and the means for discharing cooling air reduces its flow.

2. An apparatus as set forth in claim 1 wherein passage means directs compressor air on said inwardly extending flange means.

3. An appparatus as set forth in claim 1 wherein said engine is a fan engine having a fan, said fan discharging fan air into the space between the inner casing and outer casing, said fan being the means for passing a flow of cooling air over said cooling fln means.

4. An appratus as set forth in claim 2 wherein said engine is a fan engine having a fan, said fan discharging fan air into the space between the inner casing and outer casing; said fan being the means for passing a flow of cooling air over said cooling fin means.

5. An apparatus as set forth in claim 1 wherein said shroud means comprises a cylindrical-like member having an inner seal face for cooperation with said outer tips of said blades, the outer surface of said cylindrical-like member has a plurality of flow paths therein, a cylindrical-like liner is fixedly positioned adjacent said flow paths forming enclosed passageways, passage means connects air from said compressor to said passageways, opening means connects the rearward ends of said passageways to a point rearwardly of said turbine blades.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2859934 *Jul 16, 1954Nov 11, 1958Havilland Engine Co LtdGas turbines
US2880574 *May 18, 1956Apr 7, 1959Curtiss Wright CorpBy-pass turbo jet engine construction
US3391904 *Nov 2, 1966Jul 9, 1968United Aircraft CorpOptimum response tip seal
US3451215 *Apr 3, 1967Jun 24, 1969Gen ElectricFluid impingement starting means
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US3836279 *Feb 23, 1973Sep 17, 1974United Aircraft CorpSeal means for blade and shroud
US3975112 *Jun 9, 1975Aug 17, 1976United Technologies CorporationApparatus for sealing a gas turbine flow path
US3981609 *Jun 2, 1975Sep 21, 1976United Technologies CorporationCoolable blade tip shroud
US3982390 *Jan 7, 1976Sep 28, 1976General Motors CorporationGas turbine engine cooling system
US4019320 *Dec 5, 1975Apr 26, 1977United Technologies CorporationExternal gas turbine engine cooling for clearance control
US4069662 *Dec 5, 1975Jan 24, 1978United Technologies CorporationClearance control for gas turbine engine
US4177004 *Oct 31, 1977Dec 4, 1979General Electric CompanyCombined turbine shroud and vane support structure
US4242042 *May 16, 1978Dec 30, 1980United Technologies CorporationTemperature control of engine case for clearance control
US4271666 *Aug 20, 1979Jun 9, 1981Avco CorporationIntegral infrared radiation suppressor for a turbofan engine
US4353679 *Jul 29, 1976Oct 12, 1982General Electric CompanyFluid-cooled element
US4485620 *Mar 3, 1982Dec 4, 1984United Technologies CorporationCoolable stator assembly for a gas turbine engine
US4487016 *Oct 1, 1980Dec 11, 1984United Technologies CorporationModulated clearance control for an axial flow rotary machine
US4553901 *Dec 21, 1983Nov 19, 1985United Technologies CorporationStator structure for a gas turbine engine
US4573866 *May 2, 1983Mar 4, 1986United Technologies CorporationSealed shroud for rotating body
US4642024 *Dec 5, 1984Feb 10, 1987United Technologies CorporationCoolable stator assembly for a rotary machine
US4643638 *Dec 21, 1983Feb 17, 1987United Technologies CorporationStator structure for supporting an outer air seal in a gas turbine engine
US4650394 *Nov 13, 1984Mar 17, 1987United Technologies CorporationCoolable seal assembly for a gas turbine engine
US4650395 *Dec 21, 1984Mar 17, 1987United Technologies CorporationCoolable seal segment for a rotary machine
US4721433 *Dec 19, 1985Jan 26, 1988United Technologies CorporationCoolable stator structure for a gas turbine engine
US4767260 *Nov 7, 1986Aug 30, 1988United Technologies CorporationFor a gas turbine engine
US5048288 *Nov 13, 1990Sep 17, 1991United Technologies CorporationCombined turbine stator cooling and turbine tip clearance control
US5123241 *Oct 10, 1990Jun 23, 1992Societe Nationale D'etude Et De Construction De Moteurs D'aviation ("S.N.E.C.M.A.")System for deforming a turbine stator housing
US5212940 *Apr 16, 1991May 25, 1993General Electric CompanyTip clearance control apparatus and method
US6089821 *May 5, 1998Jul 18, 2000Rolls-Royce PlcGas turbine engine cooling apparatus
US6899518Dec 23, 2002May 31, 2005Pratt & Whitney Canada Corp.Turbine shroud segment apparatus for reusing cooling air
US7114339 *Mar 30, 2004Oct 3, 2006United Technologies CorporationCavity on-board injection for leakage flows
US7534088Jun 19, 2006May 19, 2009United Technologies CorporationFluid injection system
US8668431 *Mar 29, 2010Mar 11, 2014United Technologies CorporationSeal clearance control on non-cowled gas turbine engines
US20110236179 *Mar 29, 2010Sep 29, 2011United Technologies CorporationSeal clearance control on non-cowled gas turbine engines
US20110262265 *Apr 7, 2011Oct 27, 2011Rolls-Royce PlcInstallation having a thermal transfer arrangement
US20110268580 *Sep 10, 2009Nov 3, 2011Roderich BrykAxially segmented guide vane mount for a gas turbine
DE2621913A1 *May 17, 1976Dec 23, 1976United Technologies CorpGasturbinentriebwerk
DE3015653A1 *Apr 21, 1980Oct 29, 1981Avco CorpLuftgekuehltes schaufelversteifungsband eines turbinenrotors mit halterungsmitteln
EP0877149A2 *Apr 29, 1998Nov 11, 1998Rolls-Royce LimitedCooling of a gas turbine engine housing
WO2004057159A1 *Nov 18, 2003Jul 8, 2004Pratt & Whitney CanadaCooling a turbine shroud segment abnd reusing the cooling air
Classifications
U.S. Classification60/806, 415/117, 415/15, 415/116
International ClassificationF01D11/18, F01D25/12
Cooperative ClassificationF01D25/12, F01D11/18, Y02T50/675, Y02T50/671
European ClassificationF01D25/12, F01D11/18