US3391904A - Optimum response tip seal - Google Patents

Optimum response tip seal Download PDF

Info

Publication number
US3391904A
US3391904A US591524A US59152466A US3391904A US 3391904 A US3391904 A US 3391904A US 591524 A US591524 A US 591524A US 59152466 A US59152466 A US 59152466A US 3391904 A US3391904 A US 3391904A
Authority
US
United States
Prior art keywords
retainer
wear strip
turbine
flange
passageway
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime
Application number
US591524A
Other languages
English (en)
Inventor
Kenneth J Albert
Harry J Young
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Technologies Corp
Original Assignee
United Aircraft Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by United Aircraft Corp filed Critical United Aircraft Corp
Priority to US591524A priority Critical patent/US3391904A/en
Priority to GB46127/67A priority patent/GB1199974A/en
Priority to DE19671601676 priority patent/DE1601676A1/de
Priority to FR1604778D priority patent/FR1604778A/fr
Application granted granted Critical
Publication of US3391904A publication Critical patent/US3391904A/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

Images

Classifications

    • 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

Definitions

  • This invention relates to a seal construction, the rate of expansion or contraction of which closely follows the rate of expansion or contraction of a rotating rotor construction.
  • This seal construction thereby provides the minimum possible blade-to-seal gap during steady-state engine operation and also eliminates transient interference during acceleration or deceleration.
  • This invention relates to gas turbine engines and more particularly to an arrangement for maintaining the turbine blade tip clearance at the most optimum level for varying flight conditions.
  • the turbine rotor In a gas turbine engine, the turbine rotor must rotate freely at all flight conditions, including transients, so therefore there must be clearance between the turbine blade tips and the surrounding turbine casing seal assemblies.
  • An important feature of this invention is an arrangement by which the rate of expansion or contraction of the turbine casing seal assemblies is made to closely follow the rate of expansion and contraction of the turbine rotor.
  • the tip clearance is excessive at steady-state conditions because a large cold clearance is necessary to avoid interference durin transient operations, and because the high steady-state operating temperature of the turbine casing seal assembly produces excessive growth of this assembly relative to the turbine rotor growth. It should be clear that a reduction in this steady-state tip clearance will produce a significant increase in turbine efliciency with a corresponding decrease in specific fuel consumption. To this extent, it is another feature of this invention to provide a minimum hot running clearance, thereby increasing the turbine efliciency.
  • FIGURE 1 is a sectional view of a turbine section of a gas turbine engine showing the device of the invention thereof.
  • FIGURE 2 is a sectional view of a turbine section of a gas turbine engine showing a second embodiment of the device of the invention.
  • FIGURE 3 is a sectional view substantially along the line 33 of FIGURE 1.
  • FIGURE 4 is a graphical presentation of the comparative expansion and response of different type turbine seal assemblies.
  • the numeral 2 indicates a turbine generally. It is understood that all portions and parts of the turbine shown in the referenced figures may be assumed to be consistent with conventionally known gas turbine engines.
  • the turbine is provided with a rotatable disc 4, the shafts and supporting bearings not being shown.
  • the disc 4 includes the usual turbine blades '6, disposed radially around the outer periphery of the disc (FIGURES 1 and 2).
  • the engine has a forward casing 8, an inner casing 10 and a rear casing 12.
  • the outer casing 8 has an outwardly directed flange 9, the inner casing 10 has an outwardly directed flange 11 and the rear casing 12 has an outwardly directed flange 13.
  • the flanges are connected together by bolt 14, thereby connecting the turbine casings together.
  • Interposed between flange 11 and flange 13 and connected thereto by bolt 14 is an inwardly directed flange 16.
  • Inner casing 10 is radially spaced from forward casing 8 forming an annular passageway for compressor bleed cooling air.
  • Flanges 11, 13 and 16 have axially aligned holes which communicate with this annular passageway and thereby allow cooling air to pass through the flanges.
  • Flange 16 supports vanes 18 which direct the combustion gases at a predetermined angle onto blades 6.
  • An inwardly directed flange 20 connected to rear casing 12 and axially spaced from the inwardly directed end of flange 13, cooperates with flange 13 in supporting annular ring 22.
  • the ring forms a chamber between the ring and rear casing 12; however, the ring 22 has openings which allow cooling air to pass through the ring and radially inward.
  • Flange 20 also supports vane mounting ring 24, this mounting ring supporting vanes 26.
  • circumferential wear strip retainer 32 Cooperating between flange 16 and vane mounting flange 24 is circumferential wear strip retainer 32.
  • the inner circumferential end which is in contact with the first gas has a plurality of axial grooves extending partially through this face to provide for circumferential thermal expansion.
  • the wear strip retainer 32 has axial directed projections 30 which are slidable within radial openings 28 contained at the inner end of flange 16. It is to be understood that these axial directed projections may be on either side of the wear strip retainer and could cooperate within radial openings contained in the vane mounting flange.
  • the radial openings 28 are arranged in such a manner that the projections are restrained circumferentially and therefore they are held concentric with the engine center line and correspondingly, the wear strip retainer is held concentric with the engine center line. Due to the loading of the gas stream, the downstream radial face 33 of the wear strip retainer abuts and is slidable against the vane mounting flange thereby effecting a seal. Slidably connected between flange 16 and the upstream radial face 35 of wear strip retainer 32 is a conventional piston ring seal 25. Since the pressure on the internal side of the piston ring seal is greater than the gas stream pressure, the piston ring seal prevents any blow by or leakage of the cooling air between flange 16 and radial face 35.
  • circumferential slotted or segmented wear strip 34 Connected between the radial inner faces 37 of wear strip retainer 32 and radially restrained by these faces is circumferential slotted or segmented wear strip 34.
  • the slots or spacing between the segments provide for circumferential thermal expansion (see FIGURE 3).
  • the circumferential wear strip forms a chamber between the wear strip 34 and axial directed inner face 40 of wear strip retainer 32.
  • An axial and circumferential first baffle 36 radially spaced from inner face 40 and radially restrained by faces 37 forms cooling air passageway 43 between first baffle 36 and inner face 40 of wear strip retainer 32.
  • the radial distance between the first baffle and inner face 40 is critical in that it determines the heat transfer flow coeificient and thereby controls the steadystate temperature and thermal response rate of the wear strip retainer 32.
  • the present embodiment includes a second axial and circumferential baflie 38 which is radially spaced from the outer axial directed face 42 of wear strip retainer 32 and forms cooling air passageway 39 therebetween. It should be clear that the radial spacing of one baffle is suflicient to obtain the desired flow coefficient and thereby maintain the wear strip retainer at the desired steady-state temperature and the desired thermal response rate.
  • the Wear strip retainer, wear strip and first bafile are connected together by pin 44 to insure that as the parts expand and contract thermally that they act as a unitary assembly.
  • FIGURE 2 illustrates a second embodiment of the invention and in this figure the numeral 102 refers to a turbine generally and includes all parts and portions which may be assumed to be consistent with conventionally known gas turbine engines.
  • the turbine is provided with a rotatable disc 104, the shafts and supporting hearings not being shown.
  • the disc 104 includes the usual turbine blades 106, isposed radially around the outer periphery of the disc.
  • the engine has a forward casing 108 and a rear casing 110, the casings being connected together at outwardly directed flanges 109 and 111 by bolt 114.
  • Casing 108 supports an inwardly directed flange 112, this flange having means for allowing cooling air to pass through it.
  • a second inwardly directed flange 116 is also supported from casing 108.
  • Vane 118 is supported and positioned between flanges 112 and 116, and vane 118 directs the combustion gases at a predetermined angle onto blades 106.
  • an inwardly directed flange 120 Supported from rear casing 110 is an inwardly directed flange 120.
  • Flange 12% supports and positions vane 122 at vane mounting platform 124.
  • Cooperating between flange 112 and vane mounting platform 124 is circumferential wear strip retainer 126.
  • the inner circumferential face 127 which is in contact with the hot gases has a plurality of axial grooves extending partially through the face to provide for circumferential thermal expansion.
  • Wear strip retainer 126 has axial directed projections 128 which are slidable within radial openings 130 contained at the inner end of flange 112. As in the embodiment shown in FIG- URE 1, these openings are arranged such that the wear strip retainer is held concentric with the engine center line.
  • seal ring 140 Connected between inner axial face 132 of wear strip retainer 126 and outer axial face 134 of wear strip retainer is slotted or segmented circumferential wear strip 136, the slots or spacing between the segments providing for circumferential thermal expansion. Positioned between and abutting against outer radial face 138 of wear strip 136 and vane mounting platform 124 is seal ring 140.
  • the circumferential wear strip 136 forms an annular chamber between axial inner face 142 of wear strip retainer 126 and axial surface 144 of wear strip 136.
  • This chamber is divided into two chambers by an axial and circumferential first baffle 146, the chamber between axial inner face 142 and the baflie being a cooling air passageway 148 and the chamber between the baffle and axial surface 144 being a dead air chamber.
  • the embodiment in FIGURE 2 includes a second axial and circumferential baflie 150 but as noted previously one bafiie is adequate since the flow coeflicient can be obtained by the radial spacing between first bafiie 145 and inner face 142.
  • Pin 156 secures wear strip retainer ring 126, wear strip 136 and first baflie 146 together to insure that they act as a unitary assembly.
  • Cooling air bled from the compressor is supplied to the cooling air passageways, that is,
  • FIGURE 4 graphically illustrates the turbine blade and disc growth and how the turbine seal assembly responds.
  • FIGURE 4 also provides a comparison between the invention herein and a conventional rub strip assembly and is based on specific gas turbine engine design. It can be seen from this figure that for a conventional rub strip in this particular engine design, minimum clearances occur at about 30 seconds from the start of deceleration from sea level takeoff to idle steady state. As a result, steady-state clearances at cruise and sea level take-off powers are in the order of .086 inch radially and .100 inch radially respectively. This results in a serious turbine efficiency penalty.
  • the rub strip assembly described herein, since it operates at lower temperatures and its response is matched closely to the turbine rotor, it is estimated that steady-state clearances can be reduced to as low as .020 inch radially. This will result in an improvement of the turbine efliciency by as much as 2.2 percent.
  • a generally circumferential wear strip retainer said retainer having means for positioning and holding said retainer within a turbine casing and for providing for radial expansion within said casing, said retainer having a pluraiity of passageways, said passageways directing air around and through said retainer into the engine gas stream;
  • a second bafiie extends axially and circumferentially around the outer diameter of said wear strip retainer, said second baffle thereby forming a passageway between said wear strip retainer and said second bar.
  • the flow coefficient through said passageway being controlled by the radial spacing between the second baffle and the wear strip retainer, the flow coeflicient of this passageway contributing to and in some instances soiely controhing the thermal expansion and contraction response of said retainer, said passageway communicating with said passageways insaid wear strip retainer.
  • said wear strip retainer has a plurality of axial grooves, said grooves extending partially through said ring so as to provide for the circumferential thermal expansion of said retainer;
  • said wear strip has a plurality of axial slots to provide for the circumferential thermal expansion of said wear strip.
  • said wear strip is a plurality of segments, said segments thereby providing for circumferential thermal expansion of said wear strip.
  • a turbine blade seal as- 10 sembly surrounding and radially spaced from the blade tips of a turbine rotor assembly comprising:
  • said retainer having means cooperating slidably within said radial openings, thereby maintaining said retainer concentric with the engine axis and providing for radial movement of said retainer in said casing, and said retainer having a plurality of passageways therethrough for directing air through and around said retainer and into the engine gas stream,
  • said wear strip retainer has axial grooves extending partially through the portion of said wear strip retainer that is in contact with the engine gas stream, and
  • said wear strip has a plurality of axial slots thereby providing for the circumferential thermal expansion of said wear strip.
  • said first baffle divides said annular chamber between said wear strip retainer and said wear strip into two passageway between said battle and said wear strip retainer, the flow coeflicient through said passageway being controlled by the radial spacing between the first baflie and the wear strip retainer, the flow coefiicient controlling the thermal expansion and contraction of said retainer, said passageway communieating with said retainer passageways, and
  • the chamber between said first baffle and said wear strip retainer being a passageway in communication with the other flow passageways within said wear strip retainer, and the chamber between said first bafile and said wear strip being a dead air chamber, thereby being a thermal shield and assisting in maintaining said wear strip retainer means for connecting said retainer, said wear strip and said first baflle thereby insuring that they move radially as a unitary assembly.
  • said wear strip retainer has axial grooves extending partially through the portion of said wear strip retainer that is exposed to the engine gas stream
  • said wear strip has a plurality of axial slots thereby providing for the circumferential thermal expansion of said wear strip.
  • a turbine blade seal assembly comprising:
  • said flange having radial slots a generally circumferential wear strip retainer, said retainer having axially directed lugs, said lugs being slidable within said slots in retainer, said retainer having a plurality of passageways, said passageways directing air through and around said retainer into at a temperature lower than said wear strip.
  • a second baflle extending axially and circumferentially around the outer diameter of said wear strip retainer, said second baflle forming a passageway between said wear strip retainer and said second bafiie, the flow coefficient through said passageway being controlled by the radial spacing between the second bafii'e and the wear strip retainer, the flow coeflicient of this passageway contributing to and in some instances solely controlling the thermal expansion and contraction response of the retainer.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US591524A 1966-11-02 1966-11-02 Optimum response tip seal Expired - Lifetime US3391904A (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
US591524A US3391904A (en) 1966-11-02 1966-11-02 Optimum response tip seal
GB46127/67A GB1199974A (en) 1966-11-02 1967-10-09 Turbine Blade Seal Assembly
DE19671601676 DE1601676A1 (de) 1966-11-02 1967-10-10 Optimal ansprechende Turbinenblattspitzendichtung
FR1604778D FR1604778A (de) 1966-11-02 1967-10-27

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US591524A US3391904A (en) 1966-11-02 1966-11-02 Optimum response tip seal

Publications (1)

Publication Number Publication Date
US3391904A true US3391904A (en) 1968-07-09

Family

ID=24366816

Family Applications (1)

Application Number Title Priority Date Filing Date
US591524A Expired - Lifetime US3391904A (en) 1966-11-02 1966-11-02 Optimum response tip seal

Country Status (4)

Country Link
US (1) US3391904A (de)
DE (1) DE1601676A1 (de)
FR (1) FR1604778A (de)
GB (1) GB1199974A (de)

Cited By (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3529906A (en) * 1968-10-30 1970-09-22 Westinghouse Electric Corp Static seal structure
US3536365A (en) * 1968-05-31 1970-10-27 Caterpillar Tractor Co Sealing means for high speed shafts
US3583824A (en) * 1969-10-02 1971-06-08 Gen Electric Temperature controlled shroud and shroud support
US3709637A (en) * 1970-08-14 1973-01-09 Secr Defence Gas turbine engines
US3736751A (en) * 1970-05-30 1973-06-05 Secr Defence Gap control apparatus
US3742705A (en) * 1970-12-28 1973-07-03 United Aircraft Corp Thermal response shroud for rotating body
US3765791A (en) * 1970-11-07 1973-10-16 Motoren Turbinen Union Turbine nozzle support
US3807891A (en) * 1972-09-15 1974-04-30 United Aircraft Corp Thermal response turbine shroud
US3893786A (en) * 1973-06-07 1975-07-08 Ford Motor Co Air cooled shroud for a gas turbine engine
DE2855157A1 (de) * 1977-12-21 1979-06-28 United Technologies Corp Dichtungsspaltsteuerverfahren und -system fuer ein gasturbinentriebwerk
FR2416345A1 (fr) * 1978-01-31 1979-08-31 Snecma Dispositif de refroidissement par impact des segments d'etancheite de turbine d'un turboreacteur
WO1979001008A1 (en) * 1978-05-01 1979-11-29 Caterpillar Tractor Co A turbine shroud assembly
US4230439A (en) * 1978-07-17 1980-10-28 General Electric Company Air delivery system for regulating thermal growth
US4230436A (en) * 1978-07-17 1980-10-28 General Electric Company Rotor/shroud clearance control system
US4247248A (en) * 1978-12-20 1981-01-27 United Technologies Corporation Outer air seal support structure for gas turbine engine
US4307993A (en) * 1980-02-25 1981-12-29 Avco Corporation Air-cooled cylinder with piston ring labyrinth
US4439982A (en) * 1979-02-28 1984-04-03 Mtu Motoren-Und Turbinen-Union Munchen Gmbh Arrangement for maintaining clearances between a turbine rotor and casing
FR2552165A1 (fr) * 1983-09-16 1985-03-22 Mtu Muenchen Gmbh Installation pour le blocage axial et peripherique de composants de carter statiques de machines a fluide
DE3446385A1 (de) * 1983-12-21 1985-07-04 United Technologies Corp., Hartford, Conn. Statoraufbau zum abstuetzen einer aeusseren luftdichtung in einem gasturbinen-triebwerk
US4573866A (en) * 1983-05-02 1986-03-04 United Technologies Corporation Sealed shroud for rotating body
US4596116A (en) * 1983-02-10 1986-06-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Sealing ring for a turbine rotor of a turbo machine and turbo machine installations provided with such rings
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
US4650394A (en) * 1984-11-13 1987-03-17 United Technologies Corporation Coolable seal assembly for a gas turbine engine
US4668164A (en) * 1984-12-21 1987-05-26 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US4767260A (en) * 1986-11-07 1988-08-30 United Technologies Corporation Stator vane platform cooling means
US5188506A (en) * 1991-08-28 1993-02-23 General Electric Company Apparatus and method for preventing leakage of cooling air in a shroud assembly of a gas turbine engine
US5219268A (en) * 1992-03-06 1993-06-15 General Electric Company Gas turbine engine case thermal control flange
US5601402A (en) * 1986-06-06 1997-02-11 The United States Of America As Represented By The Secretary Of The Air Force Turbo machine shroud-to-rotor blade dynamic clearance control
US20050265827A1 (en) * 2002-09-09 2005-12-01 Florida Turbine Technologies, Inc. Passive clearance control
JP2007085346A (ja) * 2005-09-23 2007-04-05 Snecma ガスタービン内の間隔調整装置
US20080050225A1 (en) * 2005-03-24 2008-02-28 Alstom Technology Ltd Heat accumulation segment
US20080050224A1 (en) * 2005-03-24 2008-02-28 Alstom Technology Ltd Heat accumulation segment
US20130031914A1 (en) * 2011-08-02 2013-02-07 Ching-Pang Lee Two stage serial impingement cooling for isogrid structures
US20200291803A1 (en) * 2019-03-13 2020-09-17 United Technologies Corporation Boas carrier with dovetail attachments
CN115070662A (zh) * 2021-03-15 2022-09-20 中国航发商用航空发动机有限责任公司 涡轮盘组件的叶片挡环安装工具

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1484936A (en) * 1974-12-07 1977-09-08 Rolls Royce Gas turbine engines
GB1605255A (en) * 1975-12-02 1986-08-13 Rolls Royce Clearance control apparatus for bladed fluid flow machine
GB2119452A (en) * 1982-04-27 1983-11-16 Rolls Royce Shroud assemblies for axial flow turbomachine rotors
FR2803871B1 (fr) * 2000-01-13 2002-06-07 Snecma Moteurs Agencement de reglage de diametre d'un stator de turbine a gaz

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2858104A (en) * 1954-02-04 1958-10-28 A V Roe Canada Ltd Adjustable gas turbine shroud ring segments
US2859934A (en) * 1953-07-29 1958-11-11 Havilland Engine Co Ltd Gas turbines
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
US2962256A (en) * 1956-03-28 1960-11-29 Napier & Son Ltd Turbine blade shroud rings
US2994472A (en) * 1958-12-29 1961-08-01 Gen Electric Tip clearance control system for turbomachines
US3056583A (en) * 1960-11-10 1962-10-02 Gen Electric Retaining means for turbine shrouds and nozzle diaphragms of turbine engines
US3092393A (en) * 1958-01-20 1963-06-04 Rolls Royce Labyrinth seals
US3227418A (en) * 1963-11-04 1966-01-04 Gen Electric Variable clearance seal
GB1020900A (en) * 1961-11-28 1966-02-23 Licentia Gmbh A seal between the rotor blades and the casing of axial-flow turbo-machines
US3243158A (en) * 1964-01-15 1966-03-29 United Aircraft Corp Turbine construction

Patent Citations (10)

* 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
US2858104A (en) * 1954-02-04 1958-10-28 A V Roe Canada Ltd Adjustable gas turbine shroud ring segments
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
US2962256A (en) * 1956-03-28 1960-11-29 Napier & Son Ltd Turbine blade shroud rings
US3092393A (en) * 1958-01-20 1963-06-04 Rolls Royce Labyrinth seals
US2994472A (en) * 1958-12-29 1961-08-01 Gen Electric Tip clearance control system for turbomachines
US3056583A (en) * 1960-11-10 1962-10-02 Gen Electric Retaining means for turbine shrouds and nozzle diaphragms of turbine engines
GB1020900A (en) * 1961-11-28 1966-02-23 Licentia Gmbh A seal between the rotor blades and the casing of axial-flow turbo-machines
US3227418A (en) * 1963-11-04 1966-01-04 Gen Electric Variable clearance seal
US3243158A (en) * 1964-01-15 1966-03-29 United Aircraft Corp Turbine construction

Cited By (43)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3536365A (en) * 1968-05-31 1970-10-27 Caterpillar Tractor Co Sealing means for high speed shafts
US3529906A (en) * 1968-10-30 1970-09-22 Westinghouse Electric Corp Static seal structure
US3583824A (en) * 1969-10-02 1971-06-08 Gen Electric Temperature controlled shroud and shroud support
US3736751A (en) * 1970-05-30 1973-06-05 Secr Defence Gap control apparatus
US3709637A (en) * 1970-08-14 1973-01-09 Secr Defence Gas turbine engines
US3765791A (en) * 1970-11-07 1973-10-16 Motoren Turbinen Union Turbine nozzle support
US3742705A (en) * 1970-12-28 1973-07-03 United Aircraft Corp Thermal response shroud for rotating body
US3807891A (en) * 1972-09-15 1974-04-30 United Aircraft Corp Thermal response turbine shroud
US3893786A (en) * 1973-06-07 1975-07-08 Ford Motor Co Air cooled shroud for a gas turbine engine
DE2855157A1 (de) * 1977-12-21 1979-06-28 United Technologies Corp Dichtungsspaltsteuerverfahren und -system fuer ein gasturbinentriebwerk
FR2416345A1 (fr) * 1978-01-31 1979-08-31 Snecma Dispositif de refroidissement par impact des segments d'etancheite de turbine d'un turboreacteur
WO1979001008A1 (en) * 1978-05-01 1979-11-29 Caterpillar Tractor Co A turbine shroud assembly
US4251185A (en) * 1978-05-01 1981-02-17 Caterpillar Tractor Co. Expansion control ring for a turbine shroud assembly
US4230439A (en) * 1978-07-17 1980-10-28 General Electric Company Air delivery system for regulating thermal growth
US4230436A (en) * 1978-07-17 1980-10-28 General Electric Company Rotor/shroud clearance control system
US4247248A (en) * 1978-12-20 1981-01-27 United Technologies Corporation Outer air seal support structure for gas turbine engine
US4439982A (en) * 1979-02-28 1984-04-03 Mtu Motoren-Und Turbinen-Union Munchen Gmbh Arrangement for maintaining clearances between a turbine rotor and casing
US4307993A (en) * 1980-02-25 1981-12-29 Avco Corporation Air-cooled cylinder with piston ring labyrinth
US4596116A (en) * 1983-02-10 1986-06-24 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "S.N.E.C.M.A." Sealing ring for a turbine rotor of a turbo machine and turbo machine installations provided with such rings
US4573866A (en) * 1983-05-02 1986-03-04 United Technologies Corporation Sealed shroud for rotating body
FR2552165A1 (fr) * 1983-09-16 1985-03-22 Mtu Muenchen Gmbh Installation pour le blocage axial et peripherique de composants de carter statiques de machines a fluide
DE3446385A1 (de) * 1983-12-21 1985-07-04 United Technologies Corp., Hartford, Conn. Statoraufbau zum abstuetzen einer aeusseren luftdichtung in einem gasturbinen-triebwerk
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
US4668164A (en) * 1984-12-21 1987-05-26 United Technologies Corporation Coolable stator assembly for a gas turbine engine
US5601402A (en) * 1986-06-06 1997-02-11 The United States Of America As Represented By The Secretary Of The Air Force Turbo machine shroud-to-rotor blade dynamic clearance control
US4767260A (en) * 1986-11-07 1988-08-30 United Technologies Corporation Stator vane platform cooling means
US5188506A (en) * 1991-08-28 1993-02-23 General Electric Company Apparatus and method for preventing leakage of cooling air in a shroud assembly of a gas turbine engine
US5219268A (en) * 1992-03-06 1993-06-15 General Electric Company Gas turbine engine case thermal control flange
US20050265827A1 (en) * 2002-09-09 2005-12-01 Florida Turbine Technologies, Inc. Passive clearance control
US7210899B2 (en) 2002-09-09 2007-05-01 Wilson Jr Jack W Passive clearance control
US20080050225A1 (en) * 2005-03-24 2008-02-28 Alstom Technology Ltd Heat accumulation segment
US20080050224A1 (en) * 2005-03-24 2008-02-28 Alstom Technology Ltd Heat accumulation segment
US7658593B2 (en) * 2005-03-24 2010-02-09 Alstom Technology Ltd Heat accumulation segment
US7665958B2 (en) * 2005-03-24 2010-02-23 Alstom Technology Ltd. Heat accumulation segment
JP2007085346A (ja) * 2005-09-23 2007-04-05 Snecma ガスタービン内の間隔調整装置
US20130031914A1 (en) * 2011-08-02 2013-02-07 Ching-Pang Lee Two stage serial impingement cooling for isogrid structures
US8826668B2 (en) * 2011-08-02 2014-09-09 Siemens Energy, Inc. Two stage serial impingement cooling for isogrid structures
US20200291803A1 (en) * 2019-03-13 2020-09-17 United Technologies Corporation Boas carrier with dovetail attachments
US11761343B2 (en) * 2019-03-13 2023-09-19 Rtx Corporation BOAS carrier with dovetail attachments
CN115070662A (zh) * 2021-03-15 2022-09-20 中国航发商用航空发动机有限责任公司 涡轮盘组件的叶片挡环安装工具
CN115070662B (zh) * 2021-03-15 2024-01-12 中国航发商用航空发动机有限责任公司 涡轮盘组件的叶片挡环安装工具

Also Published As

Publication number Publication date
GB1199974A (en) 1970-07-22
FR1604778A (de) 1972-01-31
DE1601676A1 (de) 1970-08-06

Similar Documents

Publication Publication Date Title
US3391904A (en) Optimum response tip seal
US4708588A (en) Turbine cooling air supply system
US4425079A (en) Air sealing for turbomachines
US4013376A (en) Coolable blade tip shroud
US10364706B2 (en) Meter plate for blade outer air seal
US3312448A (en) Seal arrangement for preventing leakage of lubricant in gas turbine engines
US3742705A (en) Thermal response shroud for rotating body
US5022817A (en) Thermostatic control of turbine cooling air
US3730640A (en) Seal ring for gas turbine
US4309145A (en) Cooling air seal
US4242042A (en) Temperature control of engine case for clearance control
EP1630385B1 (de) Verfahren und Vorrichtung zur Aufrechterhaltung des Schaufelspitzenspiels einer Rotoranordnung
US3286461A (en) Turbine starter and cooling
US4326835A (en) Blade platform seal for ceramic/metal rotor assembly
US4311432A (en) Radial seal
US2951340A (en) Gas turbine with control mechanism for turbine cooling air
US3986720A (en) Turbine shroud structure
US3425665A (en) Gas turbine rotor blade shroud
JP3844609B2 (ja) 冷却されるタービンのディストリビュータのブレード
EP2375005B1 (de) Verfahren zur steuerung des schaufelspitzendichtspalts einer turbine
JPS6363721B2 (de)
US4702670A (en) Gas turbine engines
US4804310A (en) Clearance control apparatus for a bladed fluid flow machine
US3966352A (en) Variable area turbine
US3437313A (en) Gas turbine blade cooling