US3824031A - Turbine casing for a gas turbine engine - Google Patents

Turbine casing for a gas turbine engine Download PDF

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Publication number
US3824031A
US3824031A US00319763A US31976372A US3824031A US 3824031 A US3824031 A US 3824031A US 00319763 A US00319763 A US 00319763A US 31976372 A US31976372 A US 31976372A US 3824031 A US3824031 A US 3824031A
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US
United States
Prior art keywords
casing
struts
turbine
support
segments
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
US00319763A
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English (en)
Inventor
R Gilbert
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.)
Rolls Royce PLC
Original Assignee
Rolls Royce 1971 Ltd
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 Rolls Royce 1971 Ltd filed Critical Rolls Royce 1971 Ltd
Application granted granted Critical
Publication of US3824031A publication Critical patent/US3824031A/en
Anticipated expiration legal-status Critical
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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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
    • F01D25/162—Bearing supports
    • 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

  • ABSTRACT [30] Foreign Application Priority Data A b f b i tur me casmg or a gas tur me comprlses a casmg Janv 12, 1972 Great Brltaln 1365/72 defining the boundary of part of the turbine flow duct [52] U 5 Cl 415/135 415/138 60/39 32 and formed from a plurality of segments, each seg- [51] hit Cl 0 d 25 b 25/'26 ment being mounted from fixed structure by a plural- [58] Field 's 60/39 ity of radially extending struts.
  • At least some of the R struts are arranged to deform to allow relative thermal expansion between the hot segments and the supporting casing with at least one of the struts interposed be- [56] E S;?, EE tween the deformable struts being relatively rigid.
  • This invention relates to a turbine casing for a gas turbine engine.
  • the present invention provides a turbine casing which allows that part of the casing which forms the outer wall of the main flow duct to remain reasonably concentric and hence enables control of the leakage past the turbine rotor tips.
  • a turbine casing for a gas turbine engine comprises a plurality of separate arcuate segments which together form an annular casing defining the boundary of the gas flow duct around a turbine rotor stage, each segment being mounted from an outer supporting casing by means of a plurality of circumferentially spaced apart radially extending struts at least the majority of which are adapted to deform to allow relative thermal expansion between the segments and the supporting casing and with at least one strut interposed between the majority being relatively rigid.
  • strut Preferably there is a central, relatively rigid strut for each said segment and two further resilient struts equispaced from the central strut.
  • Said struts preferably extend in a plane substantially parallel to the turbine axis.
  • the supporting casing supports the shaft by means of two bearing panels between which the supporting casing extends.
  • FIG. 1 is a partly broken away view of a gas turbine engine whose turbine casing is in accordance with the invention
  • FIG. 2 is an enlarged section through the turbine of FIG. 1, and
  • FIG. 3 is a section on the line 3-3 of FIG. 2.
  • FIG. 1 there is shown a gas turbine engine comprising an outer casing within which are mounted in flow series a low pressure compressor 11, high pressure compressor 12, combustion section 13, high pressure turbine 14, low pressure turbine and an exhaust nozzle 16.
  • the low pressure compressor and turbine and the high pressure compressor and turbine are respectively interconnected by low pressure and high pressure shafts l7 and 18 respectively.
  • Operation of the engine is conventional in that the compressors 11 and 12 compress the incoming air which is then combusted in the combustion section 13.
  • the resulting gases exhaust through the high pressure and low pressure turbines l4 and 15 hence driving theirrespective compressors and the exhaust gases then flow through the nozzle 16 to provide propulsive thrust.
  • the engine concerned has a very high pressure ratio and consequently it is necessary to provide adequate sealing between the turbine blade tips and the surrounding casing and this is carried out by the construction shown in more detail in FIGS. 2 and 3.
  • FIG. 2 there is shown the downstream extremity 19 of the combustion chamber of the engine which is attached to a nozzle guide vane 20 .
  • a bearing panel is provided which comprises a plurality of separate struts 21 each passing through the hollow interior of one of the vanes 20 from the inner structural casing 22 of the engine.
  • the struts 21 support at their inner extremities a roller bearing 23 which in turn carries the shaft 18.
  • the nozzle guide vanes 20 are carried on a flange 24 which is bolted to the casing 22 at 25 and which also carries a further supporting cylinder or casing 26 which defines a portion of the structural casing 22.
  • the supporting casing 26 is formed on its inner surface with a plurality of supporting struts 27.
  • the struts are resilient and relatively thin with the exception of the central strut 27 of each set and as can be seen from FIGS. 2 and 3 they extend substantiallyradially and axially of the turbine. Groups of three of the struts 27 each support an arcuate segment 28, a plurality of segments 28 together form an annular casing, generally designated at 28 and which defines an outer boundary of the gas flow duct. Spaced slightly within the segments 28 are the rotor blades 29 of the turbine 14, the blades 29 being carried on a disc 30 which is in turn carried from the shaft 18.
  • Casing 26 is also provided at its extremity distant from the flange 24 with a further flange 31 which forms the forward support of the stationary guide vanes 32, which are also supported at their downstream end from the structural casing 22.
  • the low pressure turbine 15 comprising a plurality of rotor blades 33 mounted from a disc 34 which is carried from the shaft 17.
  • the tips of the blades 33 lie just inside an annular boundary casing generally designated at 36', formed in a similar way to that of the high pressure turbine.
  • the structural casing 22 directly carries a plurality of struts 35 similar to the struts 27, and once again .groups of three of these struts each support an arcuate segment 36 of the boundary casing 36.
  • the segments 36 form the annular casing 36 which is the outer boundary of the flow duct around the turbine blades 33.
  • vanes 37 Downstream of the turbine blades 33 are mounted a plurality of vanes 37 which are supported from the casing 22. These vanes are hollow and a plurality of struts 38 pass through the hollow centres of the vanes.
  • the struts 38 extend from the casing 22 to form a bearing panel which supports a roller bearing 39 which supports an extension of the shaft 17 and hence the low pressure turbine 15.
  • the casing 22 and the shafts 17 and 18 are maintained in fixed relationship with one another by the bearing panels 21 and 38, and the casing 22 and/or 26 is maintained relatively cool.
  • the segments 28 and 36 mounted either directly or indirectly from the casing 22 by their respective struts are hence held completely concentric overall with the shafts 17 and 18. Any relative expansion between the hot segments and the cold casing is allowed by deformation of the outer two resilient struts holding each segment. Although such expansion may produce distortion of each segment, the overall casing will depart very little from concentricity.
  • a turbine casing assembly for a gas turbine engine comprising:
  • a plurality of separate arcuate segments spaced radially inwardly of the annular support casing to define an annular boundary casing around at least one stage of a turbine rotor; and means to support each of said arcuate segments from said annular support casing whereby there can be circumferential thermal expansion of said segments and said boundary casing defined thereby relative to said annular support casing, said last-mentioned means including a plurality of circumferentially spaced apart radially extending struts greater than two for each arcuate segment to support the same from said annular support casing, at least one of said plurality of struts is relatively rigid and is interposed between the remaining of said plurality of struts with the same number ofthe remaining struts being on either side of the at least one relatively rigid strut, the remaining of said plurality of struts being resilient and deformable out of their radial alignment to allow for the thermal expansion.
  • each of said plurality of struts extends longitudinally and substantially parallel to the axis of the turbine.
  • each of said plurality of struts for each arcuate segment is an odd number of struts with the central strut being relatively rigid and the remaining struts being resilient and deformable.
  • a turbine casing assembly as claimed in claim 1 and comprising a turbine shaft supported from said support casing, and two bearing panels between which the casing extends and which support the shaft or shafts.
  • each of said plurality of struts for each arcuate segment includes three struts with the central strut being rigid.
  • a turbine casing assembly as claimed in claim 1 including a further annular support casing spaced radially outwardly from said first-mentioned annular support casing for supporting the same.
  • each strut of said plurality of struts for each of said arcuate segments is structurally integral with the respective arcuate segment and said support casing.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Control Of Turbines (AREA)
US00319763A 1972-01-12 1972-12-29 Turbine casing for a gas turbine engine Expired - Lifetime US3824031A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB136572 1972-01-12

Publications (1)

Publication Number Publication Date
US3824031A true US3824031A (en) 1974-07-16

Family

ID=9720760

Family Applications (1)

Application Number Title Priority Date Filing Date
US00319763A Expired - Lifetime US3824031A (en) 1972-01-12 1972-12-29 Turbine casing for a gas turbine engine

Country Status (6)

Country Link
US (1) US3824031A (de)
JP (1) JPS549247B2 (de)
DE (1) DE2300354C3 (de)
FR (1) FR2167837B1 (de)
GB (1) GB1335145A (de)
IT (1) IT973079B (de)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4248569A (en) * 1978-11-13 1981-02-03 General Motors Corporation Stator mounting
US4762462A (en) * 1986-11-26 1988-08-09 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Housing for an axial compressor
US4987736A (en) * 1988-12-14 1991-01-29 General Electric Company Lightweight gas turbine engine frame with free-floating heat shield
US4989406A (en) * 1988-12-29 1991-02-05 General Electric Company Turbine engine assembly with aft mounted outlet guide vanes
GB2388407A (en) * 2002-05-10 2003-11-12 Rolls Royce Plc Gas turbine blade tip clearance control structure
US20170284225A1 (en) * 2014-08-25 2017-10-05 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine exhaust member, and exhaust chamber maintenance method
US20200072070A1 (en) * 2018-09-05 2020-03-05 United Technologies Corporation Unified boas support and vane platform
US10830097B2 (en) 2016-02-04 2020-11-10 General Electric Company Engine casing with internal coolant flow patterns

Families Citing this family (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE2907748A1 (de) * 1979-02-28 1980-09-04 Motoren Turbinen Union Einrichtung zur minimierung und konstanthaltung der bei axialturbinen vorhandenen schaufelspitzenspiele, insbesondere fuer gasturbinentriebwerke
DE2907749C2 (de) * 1979-02-28 1985-04-25 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Einrichtung zur Minimierung von Konstanthaltung des bei Axialturbinen von Gasturbinentriebwerken vorhandenen Schaufelspitzenspiels
GB2051962B (en) * 1979-06-30 1982-12-15 Rolls Royce Turbine shroud ring support
DE3018621C2 (de) * 1980-05-16 1982-06-03 MTU Motoren- und Turbinen-Union München GmbH, 8000 München Außengehäuse für Axialverdichter oder -turbinen von Strömungsmaschinen, insbesondere Gasturbinentriebwerken
FR2577282B1 (fr) * 1985-02-13 1987-04-17 Snecma Carter de turbomachine associe a un dispositif pour ajuster le jeu entre rotor et stator
US5476363A (en) * 1993-10-15 1995-12-19 Charles E. Sohl Method and apparatus for reducing stress on the tips of turbine or compressor blades

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US841650A (en) * 1903-04-29 1907-01-15 Gen Electric Means for decreasing leakage in turbines.
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
FR1227668A (fr) * 1958-06-16 1960-08-22 Gen Motors Corp Compresseur à écoulement axial

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US841650A (en) * 1903-04-29 1907-01-15 Gen Electric Means for decreasing leakage in turbines.
US2863634A (en) * 1954-12-16 1958-12-09 Napier & Son Ltd Shroud ring construction for turbines and compressors
FR1227668A (fr) * 1958-06-16 1960-08-22 Gen Motors Corp Compresseur à écoulement axial

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4248569A (en) * 1978-11-13 1981-02-03 General Motors Corporation Stator mounting
US4762462A (en) * 1986-11-26 1988-08-09 Societe Nationale D'etude Et De Construction De Moteurs D'aviation (Snecma) Housing for an axial compressor
US4987736A (en) * 1988-12-14 1991-01-29 General Electric Company Lightweight gas turbine engine frame with free-floating heat shield
US4989406A (en) * 1988-12-29 1991-02-05 General Electric Company Turbine engine assembly with aft mounted outlet guide vanes
GB2388407A (en) * 2002-05-10 2003-11-12 Rolls Royce Plc Gas turbine blade tip clearance control structure
US20040018084A1 (en) * 2002-05-10 2004-01-29 Halliwell Mark A. Gas turbine blade tip clearance control structure
US6863495B2 (en) 2002-05-10 2005-03-08 Rolls-Royce Plc Gas turbine blade tip clearance control structure
GB2388407B (en) * 2002-05-10 2005-10-26 Rolls Royce Plc Gas turbine blade tip clearance control structure
US20170284225A1 (en) * 2014-08-25 2017-10-05 Mitsubishi Hitachi Power Systems, Ltd. Gas turbine exhaust member, and exhaust chamber maintenance method
US10865658B2 (en) * 2014-08-25 2020-12-15 Mitsubishi Power, Ltd. Gas turbine exhaust member, and exhaust chamber maintenance method
US10830097B2 (en) 2016-02-04 2020-11-10 General Electric Company Engine casing with internal coolant flow patterns
US20200072070A1 (en) * 2018-09-05 2020-03-05 United Technologies Corporation Unified boas support and vane platform

Also Published As

Publication number Publication date
JPS549247B2 (de) 1979-04-23
GB1335145A (en) 1973-10-24
DE2300354A1 (de) 1973-07-26
IT973079B (it) 1974-06-10
FR2167837A1 (de) 1973-08-24
DE2300354B2 (de) 1974-06-20
JPS4880906A (de) 1973-10-30
DE2300354C3 (de) 1975-02-06
FR2167837B1 (de) 1977-04-22

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