US20140119887A1 - Fluid-cooled seal arrangement for a gas turbine engine - Google Patents

Fluid-cooled seal arrangement for a gas turbine engine Download PDF

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Publication number
US20140119887A1
US20140119887A1 US13/666,591 US201213666591A US2014119887A1 US 20140119887 A1 US20140119887 A1 US 20140119887A1 US 201213666591 A US201213666591 A US 201213666591A US 2014119887 A1 US2014119887 A1 US 2014119887A1
Authority
US
United States
Prior art keywords
passage
gas turbine
turbine engine
recited
rotational component
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.)
Abandoned
Application number
US13/666,591
Other languages
English (en)
Inventor
Stephen J. Lyle
Fungayi Mutsengi
Ernest Boratgis
Santosh Ranganath
Christopher J. Larson
M. Rifat Ullah
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.)
RTX Corp
Original Assignee
United Technologies 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 Technologies Corp filed Critical United Technologies Corp
Priority to US13/666,591 priority Critical patent/US20140119887A1/en
Assigned to UNITED TECHNOLOGIES CORPORATION reassignment UNITED TECHNOLOGIES CORPORATION ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: LARSON, CHRISTOPHER J., Mutsengi, Fungayi, ULLAH, M RIFAT, BORATGIS, ERNEST, LYLE, STEPHEN J., RANGANATH, SANTOSH
Priority to EP13874213.5A priority patent/EP2914819A4/fr
Priority to PCT/US2013/068053 priority patent/WO2014120309A2/fr
Publication of US20140119887A1 publication Critical patent/US20140119887A1/en
Abandoned legal-status Critical Current

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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
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D25/18—Lubricating arrangements
    • F01D25/183—Sealing means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/06—Arrangements of bearings; Lubricating
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/24—Three-dimensional ellipsoidal
    • F05D2250/241—Three-dimensional ellipsoidal spherical
    • 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
    • Y10—TECHNICAL SUBJECTS COVERED BY FORMER USPC
    • Y10T—TECHNICAL SUBJECTS COVERED BY FORMER US CLASSIFICATION
    • Y10T29/00—Metal working
    • Y10T29/49—Method of mechanical manufacture
    • Y10T29/49316—Impeller making
    • Y10T29/4932—Turbomachine making

Definitions

  • the present disclosure relates to a gas turbine engine and, more particularly, to fluid-cooled seal arrangements therefor.
  • Certain sections of gas turbine engines may be subjected to high temperatures and pressures.
  • Some engine components may be sensitive to the harsh environment thereof and are shielded therefrom. These components are typically vented by ambient or cooling bleed-off air or have cooling oil flowing therethrough. In order to maintain cool air in the cavities housing these components, the cavities must be shielded from engine pressure and temperature differentials.
  • seal systems are positioned to prevent high temperature and pressure air from flowing downstream into the areas with lower temperature and pressure air.
  • One such seal system includes arcuate carbon material segments arranged to form a stationary carbon ring that forms a rubbing interface with a rotating seal runner.
  • the rubbing interface between the rotating seal runner and the carbon ring minimizes or prevents leakage, however if the heat generated by the rubbing interface is not adequately dissipated, the rotating seal runner may thermally distort. This may degrade performance manifested by excessive fluid leakage.
  • One approach to minimize overheating of the seal interface includes the delivery of a cooling fluid onto the underside of the rotating seal runner sprayed from a stationary nozzle positioned proximate the rotating seal runner.
  • the relative motion between the rotating seal runner and the stationary nozzle causes a uniform film of cooling fluid to be deposited on the seal runner to extract thermal energy.
  • Stationary nozzles provide a consistently even film of cooling fluid on the underside of the rotating seal runner, however their applicability on many gas turbine engines is limited by physical constraints that prevent the nozzle from being located proximate the seal runner.
  • Another approach utilizes a rotating distributor to deliver the cooling fluid onto the underside of the rotating seal runner.
  • the rotating distributor is typically affixed to the seal runner, and a steady stream of cooling fluid is delivered through a central passageway in the rotating distributor to the underside of the seal runner.
  • a series of openings in the rotating distributor dispense the cooling fluid onto the seal runner.
  • Carbon seal systems with a rotating distributor include a significant quantity of openings to deliver an even film of cooling fluid. If this design parameter is not satisfied, an uneven film of cooling fluid is distributed across the seal runner, which may cause an uneven extraction of heat.
  • carbon seal system designs with large quantities of cooling fluid dispensing openings may, under centrifugal loads, concentrate stresses which may lead to fatigue life shortfalls.
  • expensive rotor grade material may be used to meet fatigue life requirements. In other cases even rotor grade material is insufficient to meet desired fatigue life requirements.
  • a rotational component for a gas turbine engine includes at least one passage with a semi-spherical end.
  • the at least one passage is an oil communication passage.
  • the at least one passage is a blind hole.
  • a gas turbine engine includes a rotational component having at least one passage with a semi-spherical end.
  • the rotational component is a seal runner.
  • the at least one passage is an oil communication passage.
  • the at least one passage is a blind hole.
  • the oil communication passage is an outlet passage.
  • the rotational component includes a cantilevered section.
  • the cantilevered section collects and directs a lubricant toward an internal inlet passage.
  • the cantilevered section include a hook-shaped end section.
  • a method of forming a passage in a rotation component according to another disclosed non-limiting embodiment of the present disclosure includes forming a semi-spherical end at an end of the passage.
  • the method comprises drilling the passage and inserting a ball-endmill into the passage to form the end.
  • the method includes drilling the passage with a drill tool having spherical shaped flutes.
  • FIG. 1 is a schematic cross-section of a gas turbine engine
  • FIG. 2 is a partial longitudinal schematic sectional view of a bearing compartment that may be used with the gas turbine engine shown in FIG. 1 ;
  • FIG. 3 is a sectional view of a seal runner with a passage with a semi-spherical end according to one disclosed non-limiting embodiment
  • FIG. 4 is an expanded view of a RELATED ART passage.
  • FIG. 1 schematically illustrates a gas turbine engine 20 .
  • the gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22 , a compressor section 24 , a combustor section 26 and a turbine section 28 .
  • Alternative engines might include an augmentor section (not shown) among other systems or features.
  • the fan section 22 drives air along a bypass flowpath while the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26 then expansion through the turbine section 28 .
  • turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a three-spool (plus fan) engine wherein an intermediate spool includes an intermediate pressure compressor (IPC) between the LPC and HPC and an intermediate pressure turbine (IPT) between the HPT and LPT.
  • IPC intermediate pressure compressor
  • IPT intermediate pressure turbine
  • the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38 .
  • the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42 , a low pressure compressor 44 (“LPC”) and a low pressure turbine 46 (“LPT”).
  • the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30 .
  • An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
  • the high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 (“HPC”) and high pressure turbine 54 (“HPT”).
  • a combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54 .
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • Core airflow is compressed by the low pressure compressor 44 then the high pressure compressor 52 , mixed with the fuel and burned in the combustor 56 , then expanded over the high pressure turbine 54 and low pressure turbine 46 .
  • the turbines 54 , 46 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
  • the main engine shafts 40 , 50 are supported at a plurality of points by bearing structures 38 within the static structure 36 . It should be understood that various bearing structures 38 at various locations may alternatively or additionally be provided.
  • the gas turbine engine 20 is a high-bypass geared aircraft engine.
  • the gas turbine engine 20 bypass ratio is greater than about six (6:1).
  • the geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system.
  • the example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1.
  • the geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the low pressure compressor 44 and low pressure turbine 46 and render increased pressure in a fewer number of stages.
  • a pressure ratio associated with the low pressure turbine 46 is pressure measured prior to the inlet of the low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle of the gas turbine engine 20 .
  • the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the low pressure compressor 44
  • the low pressure turbine 46 has a pressure ratio that is greater than about 5 (5:1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
  • a significant amount of thrust is provided by the bypass flow path B due to the high bypass ratio.
  • the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC).
  • TSFC Thrust Specific Fuel Consumption
  • Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system.
  • the low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45.
  • Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of “T”/518.7 0.5 in which “T” represents the ambient temperature in degrees Rankine.
  • the Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
  • a bearing structure 38 A includes a bearing 58 and fluid-cooled seal arrangements 60 .
  • the fluid-cooled seal arrangements 60 A, 60 B may each be, in the disclosed non-limiting embodiment a carbon seal system to seal a “wet” zone from a “dry” zone.
  • regions or volumes containing lubricant will be referred to as a “wet” zone and a lubricant free region will be referred to as a “dry” zone.
  • the fluid-cooled seal arrangement 60 generally includes a stationary component 62 and a rotational component 64 .
  • the stationary component 62 is coupled to a rotationally fixed structure such as the static structure 36 while the rotational component 64 is mechanically connected to a rotating structure such as the outer shaft 50 . It should be appreciated, however, that any rotating structure such as a rotor hub may alternatively mount the rotational component 64 .
  • the rotational component 64 is referred to herein as a seal runner.
  • the stationary component 62 is arranged with respect to the rotational component 64 to form a rubbing interface 66 therebetween which in the disclosed non-limiting embodiment, is axially oriented with respect to the engine axis A.
  • the rubbing interface 66 may be radially oriented.
  • a first annular surface 68 is defined by the stationary component 62 and a second annular surface 70 is defined by the rotational component 64 which are maintained in rubbing contact to form a fluid tight seal at the rubbing interface 66 .
  • the lubrication system (illustrated schematically) provides cooling fluid under pressure to lubricate and cool the moving parts of the engine 20 , such as the bearing 58 and fluid-cooled seal arrangement 60 through a nozzle 72 .
  • the lubrication system discharges the fluid from the nozzle 72 with sufficient kinetic energy to spray an underside of a cantilevered section 74 of the rotational component 64 .
  • the cantilevered section 74 may include a hook-shaped end section 76 to collect and direct the lubricant toward an internal passage 78 ( FIG. 3 ).
  • the internal inlet passage 78 may be formed from one or a multiple of drill holes.
  • the internal inlet passage 78 communicates with an internal outlet passage 82 .
  • the internal outlet passage 82 may also be formed from one or a multiple of drill holes.
  • the internal outlet passage 82 includes a semi-spherical end 90 as compared to a conventional conical end E ( FIG. 7 ; RELATED ART). It should be understood that although only the internal outlet passage 82 includes the semi-spherical end 90 , any passage may benefit herefrom such as the internal inlet passage 78 .
  • the semi-spherical end 90 is applicable to any blind hole such as a drill hole and may be manufactured with, for example only, a ball-endmill type tool which is chased down a conventional drill hole or other special drill tool with spherical shaped flutes.
  • the semi-spherical end 90 reduces concentrated stress in the passages and increases fatigue life of the rotational component 64 . That is, the semi-spherical end 90 may be particularly applicable to any rotating structure such as the disclosed seal runner which is subject to high centrifugal forces and is potentially a less expensive alternative to usage of higher grade materials to meet fatigue life requirements.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US13/666,591 2012-11-01 2012-11-01 Fluid-cooled seal arrangement for a gas turbine engine Abandoned US20140119887A1 (en)

Priority Applications (3)

Application Number Priority Date Filing Date Title
US13/666,591 US20140119887A1 (en) 2012-11-01 2012-11-01 Fluid-cooled seal arrangement for a gas turbine engine
EP13874213.5A EP2914819A4 (fr) 2012-11-01 2013-11-01 Agencement de joint d'étanchéité refroidi par fluide pour moteur à turbine à gaz
PCT/US2013/068053 WO2014120309A2 (fr) 2012-11-01 2013-11-01 Agencement de joint d'étanchéité refroidi par fluide pour moteur à turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US13/666,591 US20140119887A1 (en) 2012-11-01 2012-11-01 Fluid-cooled seal arrangement for a gas turbine engine

Publications (1)

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US20140119887A1 true US20140119887A1 (en) 2014-05-01

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US (1) US20140119887A1 (fr)
EP (1) EP2914819A4 (fr)
WO (1) WO2014120309A2 (fr)

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US9631508B2 (en) 2013-06-13 2017-04-25 Pratt & Whitney Canada Corp. Internally cooled seal runner
EP3282153A1 (fr) * 2016-08-09 2018-02-14 United Technologies Corporation Caractéristiques de refroidissement de siège de joint hydrodynamique
US20180087404A1 (en) * 2016-09-23 2018-03-29 Rolls-Royce Corporation Oil distributor
US10100672B2 (en) * 2013-03-08 2018-10-16 United Technologies Corporation Fluid-cooled seal arrangement for a gas turbine engine
US20200095891A1 (en) * 2018-09-21 2020-03-26 United Technologies Corporation Hybrid wet-dry face seal seat
EP3643887A1 (fr) * 2018-10-24 2020-04-29 United Technologies Corporation Ensemble de joints d'étanchéité de moteur à turbine à gaz
US20200271013A1 (en) * 2019-02-26 2020-08-27 United Technologies Corporation Gas turbine engine including seal plate providing increased cooling adjacent contact area
EP3808943A1 (fr) * 2019-10-18 2021-04-21 Raytheon Technologies Corporation Portée de joint d'étanchéité refroidie par fluide pour ensemble joints d'équipement rotatif
US11028717B2 (en) 2017-06-26 2021-06-08 Raytheon Technologies Corporation Bearing assembly for gas turbine engines
EP3865680A1 (fr) * 2020-02-13 2021-08-18 Raytheon Technologies Corporation Joint rotatif à refroidissement par impact
US11203948B2 (en) 2019-09-06 2021-12-21 Pratt & Whitney Canada Corp. Seal runner and method
US11248492B2 (en) 2019-03-18 2022-02-15 Raytheon Technologies Corporation Seal assembly for a gas turbine engine

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US2857182A (en) * 1955-05-31 1958-10-21 Gen Motors Corp Shaft seal
US2990202A (en) * 1958-09-18 1961-06-27 United Aircraft Corp Labyrinth face seal plate
US3147013A (en) * 1961-02-10 1964-09-01 Borg Warner Dry gas seal
US4928978A (en) * 1988-04-07 1990-05-29 United Technologies Corporation Rotating shaft seal
US5284347A (en) * 1991-03-25 1994-02-08 General Electric Company Gas bearing sealing means
US5685671A (en) * 1993-11-01 1997-11-11 Smith International, Inc. Diamond or CBN fluted center cutting end mill
US5626347A (en) * 1994-09-29 1997-05-06 Alliedsignal Inc. Coning resistant face seal having a "U" shape
US6196790B1 (en) * 1998-12-17 2001-03-06 United Technologies Corporation Seal assembly for an intershaft seal in a gas turbine engine
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Cited By (20)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10100672B2 (en) * 2013-03-08 2018-10-16 United Technologies Corporation Fluid-cooled seal arrangement for a gas turbine engine
US11066955B2 (en) 2013-03-08 2021-07-20 Raytheon Technologies Corporation Fluid-cooled seal arrangement for a gas turbine engine
US9631508B2 (en) 2013-06-13 2017-04-25 Pratt & Whitney Canada Corp. Internally cooled seal runner
US10526907B2 (en) 2013-06-13 2020-01-07 Pratt & Whitney Canada Corp. Internally cooled seal runner
EP3282153A1 (fr) * 2016-08-09 2018-02-14 United Technologies Corporation Caractéristiques de refroidissement de siège de joint hydrodynamique
US20180087404A1 (en) * 2016-09-23 2018-03-29 Rolls-Royce Corporation Oil distributor
US11028717B2 (en) 2017-06-26 2021-06-08 Raytheon Technologies Corporation Bearing assembly for gas turbine engines
US20200095891A1 (en) * 2018-09-21 2020-03-26 United Technologies Corporation Hybrid wet-dry face seal seat
US10746051B2 (en) * 2018-09-21 2020-08-18 United Technologies Corporation Hybrid wet-dry face seal seat
US10774684B2 (en) * 2018-10-24 2020-09-15 Raytheon Technologies Corporation Gas turbine engine seal assemblies
US20200131935A1 (en) * 2018-10-24 2020-04-30 United Technologies Corporation Gas turbine engine seal assemblies
EP3643887A1 (fr) * 2018-10-24 2020-04-29 United Technologies Corporation Ensemble de joints d'étanchéité de moteur à turbine à gaz
US20200271013A1 (en) * 2019-02-26 2020-08-27 United Technologies Corporation Gas turbine engine including seal plate providing increased cooling adjacent contact area
US10975723B2 (en) * 2019-02-26 2021-04-13 Raytheon Technologies Corporation Gas turbine engine including seal plate providing increased cooling adjacent contact area
US11248492B2 (en) 2019-03-18 2022-02-15 Raytheon Technologies Corporation Seal assembly for a gas turbine engine
US11203948B2 (en) 2019-09-06 2021-12-21 Pratt & Whitney Canada Corp. Seal runner and method
EP3808943A1 (fr) * 2019-10-18 2021-04-21 Raytheon Technologies Corporation Portée de joint d'étanchéité refroidie par fluide pour ensemble joints d'équipement rotatif
US11193389B2 (en) * 2019-10-18 2021-12-07 Raytheon Technologies Corporation Fluid cooled seal land for rotational equipment seal assembly
EP3865680A1 (fr) * 2020-02-13 2021-08-18 Raytheon Technologies Corporation Joint rotatif à refroidissement par impact
US11441448B2 (en) 2020-02-13 2022-09-13 Raytheon Technologies Corporation Impingement cooled rotating seal

Also Published As

Publication number Publication date
EP2914819A4 (fr) 2016-01-20
WO2014120309A2 (fr) 2014-08-07
WO2014120309A3 (fr) 2014-10-16
EP2914819A2 (fr) 2015-09-09

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Free format text: ASSIGNMENT OF ASSIGNORS INTEREST;ASSIGNORS:LYLE, STEPHEN J.;MUTSENGI, FUNGAYI;BORATGIS, ERNEST;AND OTHERS;SIGNING DATES FROM 20121030 TO 20121031;REEL/FRAME:029227/0922

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