US3742705A - Thermal response shroud for rotating body - Google Patents

Thermal response shroud for rotating body Download PDF

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US3742705A
US3742705A US00101481A US3742705DA US3742705A US 3742705 A US3742705 A US 3742705A US 00101481 A US00101481 A US 00101481A US 3742705D A US3742705D A US 3742705DA US 3742705 A US3742705 A US 3742705A
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fan
shroud
engine
flow
cooling
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US00101481A
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P Sifford
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RTX Corp
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United Aircraft Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • 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
  • 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.
  • clearance at room temperature is large but the steady state clearance is kept low by increased cooling in the case and shroud support structure.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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.
  • 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,
  • cooling fin 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.
  • passage means directs compressor air on said inwardly extending flange means.
  • 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.
  • 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.
  • 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.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

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.

Description

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.

Claims (5)

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 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 fin 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.
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Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3836279A (en) * 1973-02-23 1974-09-17 United Aircraft Corp Seal means for blade and shroud
US3975112A (en) * 1975-06-09 1976-08-17 United Technologies Corporation Apparatus for sealing a gas turbine flow path
US3981609A (en) * 1975-06-02 1976-09-21 United Technologies Corporation Coolable blade tip shroud
US3982390A (en) * 1974-08-22 1976-09-28 General Motors Corporation Gas turbine engine cooling system
US4019320A (en) * 1975-12-05 1977-04-26 United Technologies Corporation External gas turbine engine cooling for clearance control
FR2333953A1 (en) * 1975-12-05 1977-07-01 United Technologies Corp CLEARANCE ADJUSTMENT DEVICE FOR A GAS TURBINE ENGINE
FR2407342A1 (en) * 1977-10-31 1979-05-25 Gen Electric TURBOMACHINE SUPPORT STRUCTURE
JPS54159516A (en) * 1978-06-05 1979-12-17 Gen Electric Shraud support apparatus
US4242042A (en) * 1978-05-16 1980-12-30 United Technologies Corporation Temperature control of engine case for clearance control
FR2468740A1 (en) * 1979-10-31 1981-05-08 Gen Electric TURBOMACHINE COMPRISING A GAME ADJUSTMENT STRUCTURE BETWEEN THE ROTOR AND THE RUBBER SURROUNDING IT
US4271666A (en) * 1979-08-20 1981-06-09 Avco Corporation Integral infrared radiation suppressor for a turbofan engine
DE3015653A1 (en) * 1979-02-09 1981-10-29 Avco Corp., 06830 Greenwich, Conn. AIR-COOLED BLADE REINFORCING TAPE OF A TURBINE ROTOR WITH BRACKETS
US4353679A (en) * 1976-07-29 1982-10-12 General Electric Company Fluid-cooled element
FR2543219A1 (en) * 1983-03-11 1984-09-28 United Technologies Corp Stator assembly which may be cooled for a gas turbine
US4485620A (en) * 1982-03-03 1984-12-04 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4487016A (en) * 1980-10-01 1984-12-11 United Technologies Corporation Modulated clearance control for an axial flow rotary machine
US4553901A (en) * 1983-12-21 1985-11-19 United Technologies Corporation Stator structure for a gas turbine engine
US4573866A (en) * 1983-05-02 1986-03-04 United Technologies Corporation Sealed shroud for rotating body
US4642024A (en) * 1984-12-05 1987-02-10 United Technologies Corporation Coolable stator assembly for a rotary machine
US4643638A (en) * 1983-12-21 1987-02-17 United Technologies Corporation Stator structure for supporting an outer air seal in a gas turbine engine
US4650394A (en) * 1984-11-13 1987-03-17 United Technologies Corporation Coolable seal assembly for a gas turbine engine
US4650395A (en) * 1984-12-21 1987-03-17 United Technologies Corporation Coolable seal segment for a rotary machine
US4721433A (en) * 1985-12-19 1988-01-26 United Technologies Corporation Coolable stator structure for a gas turbine engine
US4767260A (en) * 1986-11-07 1988-08-30 United Technologies Corporation Stator vane platform cooling means
US5048288A (en) * 1988-12-20 1991-09-17 United Technologies Corporation Combined turbine stator cooling and turbine tip clearance control
US5123241A (en) * 1989-10-11 1992-06-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation ("S.N.E.C.M.A.") System for deforming a turbine stator housing
US5212940A (en) * 1991-04-16 1993-05-25 General Electric Company Tip clearance control apparatus and method
EP0877149A3 (en) * 1997-05-07 2000-02-02 Rolls-Royce Plc Cooling of a gas turbine engine housing
RU2210672C2 (en) * 1998-09-10 2003-08-20 Алстом Device for cooling over-rotor surfaces of turbine nozzle assembly
US20040120803A1 (en) * 2002-12-23 2004-06-24 Terrence Lucas Turbine shroud segment apparatus for reusing cooling air
US20050217277A1 (en) * 2004-03-30 2005-10-06 Ioannis Alvanos Cavity on-board injection for leakage flows
US20060147299A1 (en) * 2002-11-15 2006-07-06 Piero Iacopetti Shround cooling assembly for a gas trubine
US7534088B1 (en) 2006-06-19 2009-05-19 United Technologies Corporation Fluid injection system
US20110236179A1 (en) * 2010-03-29 2011-09-29 United Technologies Corporation Seal clearance control on non-cowled gas turbine engines
US20110262265A1 (en) * 2010-04-26 2011-10-27 Rolls-Royce Plc Installation having a thermal transfer arrangement
US20110268580A1 (en) * 2008-11-05 2011-11-03 Roderich Bryk Axially segmented guide vane mount for a gas turbine
US20120247121A1 (en) * 2010-02-24 2012-10-04 Tsuyoshi Kitamura Aircraft gas turbine
US20130045089A1 (en) * 2011-08-16 2013-02-21 Joseph W. Bridges Gas turbine engine seal assembly having flow-through tube
US20140017072A1 (en) * 2012-07-16 2014-01-16 Michael G. McCaffrey Blade outer air seal with cooling features
US20200072070A1 (en) * 2018-09-05 2020-03-05 United Technologies Corporation Unified boas support and vane platform
EP3674521A1 (en) * 2018-12-27 2020-07-01 Rolls-Royce Corporation Passive blade tip clearance control system for gas turbine engine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US2880574A (en) * 1956-05-18 1959-04-07 Curtiss Wright Corp By-pass turbo jet engine construction
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3451215A (en) * 1967-04-03 1969-06-24 Gen Electric Fluid impingement starting means

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US2880574A (en) * 1956-05-18 1959-04-07 Curtiss Wright Corp By-pass turbo jet engine construction
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3451215A (en) * 1967-04-03 1969-06-24 Gen Electric Fluid impingement starting means

Cited By (56)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3836279A (en) * 1973-02-23 1974-09-17 United Aircraft Corp Seal means for blade and shroud
US3982390A (en) * 1974-08-22 1976-09-28 General Motors Corporation Gas turbine engine cooling system
US3981609A (en) * 1975-06-02 1976-09-21 United Technologies Corporation Coolable blade tip shroud
US3975112A (en) * 1975-06-09 1976-08-17 United Technologies Corporation Apparatus for sealing a gas turbine flow path
DE2621913A1 (en) * 1975-06-09 1976-12-23 United Technologies Corp GAS TURBINE ENGINE
FR2333953A1 (en) * 1975-12-05 1977-07-01 United Technologies Corp CLEARANCE ADJUSTMENT DEVICE FOR A GAS TURBINE ENGINE
US4019320A (en) * 1975-12-05 1977-04-26 United Technologies Corporation External gas turbine engine cooling for clearance control
US4069662A (en) * 1975-12-05 1978-01-24 United Technologies Corporation Clearance control for gas turbine engine
US4353679A (en) * 1976-07-29 1982-10-12 General Electric Company Fluid-cooled element
FR2407342A1 (en) * 1977-10-31 1979-05-25 Gen Electric TURBOMACHINE SUPPORT STRUCTURE
US4177004A (en) * 1977-10-31 1979-12-04 General Electric Company Combined turbine shroud and vane support structure
US4242042A (en) * 1978-05-16 1980-12-30 United Technologies Corporation Temperature control of engine case for clearance control
JPS54159516A (en) * 1978-06-05 1979-12-17 Gen Electric Shraud support apparatus
DE3015653A1 (en) * 1979-02-09 1981-10-29 Avco Corp., 06830 Greenwich, Conn. AIR-COOLED BLADE REINFORCING TAPE OF A TURBINE ROTOR WITH BRACKETS
US4271666A (en) * 1979-08-20 1981-06-09 Avco Corporation Integral infrared radiation suppressor for a turbofan engine
FR2468740A1 (en) * 1979-10-31 1981-05-08 Gen Electric TURBOMACHINE COMPRISING A GAME ADJUSTMENT STRUCTURE BETWEEN THE ROTOR AND THE RUBBER SURROUNDING IT
US4487016A (en) * 1980-10-01 1984-12-11 United Technologies Corporation Modulated clearance control for an axial flow rotary machine
US4485620A (en) * 1982-03-03 1984-12-04 United Technologies Corporation Coolable stator assembly for a gas turbine engine
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US4573866A (en) * 1983-05-02 1986-03-04 United Technologies Corporation Sealed shroud for rotating body
US4553901A (en) * 1983-12-21 1985-11-19 United Technologies Corporation Stator structure for a gas turbine engine
US4643638A (en) * 1983-12-21 1987-02-17 United Technologies Corporation Stator structure for supporting an outer air seal in a gas turbine engine
US4650394A (en) * 1984-11-13 1987-03-17 United Technologies Corporation Coolable seal assembly for a gas turbine engine
US4642024A (en) * 1984-12-05 1987-02-10 United Technologies Corporation Coolable stator assembly for a rotary machine
US4650395A (en) * 1984-12-21 1987-03-17 United Technologies Corporation Coolable seal segment for a rotary machine
US4721433A (en) * 1985-12-19 1988-01-26 United Technologies Corporation Coolable stator structure for a gas turbine engine
US4767260A (en) * 1986-11-07 1988-08-30 United Technologies Corporation Stator vane platform cooling means
US5048288A (en) * 1988-12-20 1991-09-17 United Technologies Corporation Combined turbine stator cooling and turbine tip clearance control
US5123241A (en) * 1989-10-11 1992-06-23 Societe Nationale D'etude Et De Construction De Moteurs D'aviation ("S.N.E.C.M.A.") System for deforming a turbine stator housing
US5212940A (en) * 1991-04-16 1993-05-25 General Electric Company Tip clearance control apparatus and method
EP0877149A3 (en) * 1997-05-07 2000-02-02 Rolls-Royce Plc Cooling of a gas turbine engine housing
US6089821A (en) * 1997-05-07 2000-07-18 Rolls-Royce Plc Gas turbine engine cooling apparatus
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US20060147299A1 (en) * 2002-11-15 2006-07-06 Piero Iacopetti Shround cooling assembly for a gas trubine
US6899518B2 (en) 2002-12-23 2005-05-31 Pratt & Whitney Canada Corp. Turbine shroud segment apparatus for reusing cooling air
WO2004057159A1 (en) * 2002-12-23 2004-07-08 Pratt & Whitney Canada Corp. Cooling a turbine shroud segment abnd reusing the cooling air
US20040120803A1 (en) * 2002-12-23 2004-06-24 Terrence Lucas Turbine shroud segment apparatus for reusing cooling air
US20050217277A1 (en) * 2004-03-30 2005-10-06 Ioannis Alvanos Cavity on-board injection for leakage flows
US7114339B2 (en) * 2004-03-30 2006-10-03 United Technologies Corporation Cavity on-board injection for leakage flows
US7534088B1 (en) 2006-06-19 2009-05-19 United Technologies Corporation Fluid injection system
US8870526B2 (en) * 2008-11-05 2014-10-28 Siemens Aktiengesellschaft Axially segmented guide vane mount for a gas turbine
US20110268580A1 (en) * 2008-11-05 2011-11-03 Roderich Bryk Axially segmented guide vane mount for a gas turbine
US9945250B2 (en) * 2010-02-24 2018-04-17 Mitsubishi Heavy Industries Aero Engines, Ltd. Aircraft gas turbine
US20120247121A1 (en) * 2010-02-24 2012-10-04 Tsuyoshi Kitamura Aircraft gas turbine
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US20110262265A1 (en) * 2010-04-26 2011-10-27 Rolls-Royce Plc Installation having a thermal transfer arrangement
US20130045089A1 (en) * 2011-08-16 2013-02-21 Joseph W. Bridges Gas turbine engine seal assembly having flow-through tube
US9080449B2 (en) * 2011-08-16 2015-07-14 United Technologies Corporation Gas turbine engine seal assembly having flow-through tube
US20140017072A1 (en) * 2012-07-16 2014-01-16 Michael G. McCaffrey Blade outer air seal with cooling features
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