US3709637A - Gas turbine engines - Google Patents

Gas turbine engines Download PDF

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Publication number
US3709637A
US3709637A US00169994A US3709637DA US3709637A US 3709637 A US3709637 A US 3709637A US 00169994 A US00169994 A US 00169994A US 3709637D A US3709637D A US 3709637DA US 3709637 A US3709637 A US 3709637A
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US
United States
Prior art keywords
shaft
engine
gas turbine
compressor
rotor
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
US00169994A
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English (en)
Inventor
J Petrie
T Gardiner
J Clampitt
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SECR DEFENCE
STATE FOR DEFENCE GB
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SECR DEFENCE
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Publication date
Application filed by SECR DEFENCE filed Critical SECR DEFENCE
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Publication of US3709637A publication Critical patent/US3709637A/en
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Expired - Lifetime legal-status Critical Current

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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
    • F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/026—Shaft to shaft connections
    • 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
    • F02C7/20—Mounting or supporting of plant; Accommodating heat expansion or creep
    • 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

  • PATENTEDJM 9 I973 SHEEI10F4 FIG 1 PATENTEUJAN 9197a SHEET 2 UF 4 FIGZ GAS TURBINE ENGINES The present invention relates to gas turbine engines and has particular reference to the control of sealing clearances between turbine rotors and adjacent static structure.
  • the object of the present invention is to provide a construction of gas turbine engine in which a sealing clearance between a rotor member and adjacent static structure is controlled to a low value, particularly at the engine design condition.
  • a gas turbine engine comprises a rotor member, a shaft to which the rotor member is connected and which is mounted in bearings which allow freedom for axial thermal expansion of the shaft, the rotor member being spaced from a sealing member supported in static structure at a first radial plane of the engine by an axial sealing clearance, the shaft being located in static structure at a second datum radial plane of the engine by means of a member made of a material having a lower coefficient of thermal expansion than the shaft.
  • the rotor member is a turbine rotor and the shaft drivingly connects the turbine rotor to a compressor, the member which locates the shaft is a tubular member which is located by means of a ball bearing in said static structure upstream of the compressor.
  • Part of the engine casing to which the said static structure is attached may also be made from the low expansion material to reduce the axial growth of the casmg.
  • the position of the connection between the tubular member and the shaft must be calculated from the lengths and temperatures of the shaft, the tubular member and the casing.
  • the various components are arranged to be washed by air bled from the compressor in known quantities and at known temperatures.
  • the tubular member is additionally made as thin as possible to ensure a quick response to changes in temperature.
  • FIG. 1 illustrates diagrammatically a gas turbine engine to which the invention is applicable.
  • FIG. 2 is a detailed sectional view of part of the gas turbine engine of FIG. I constructed according to this invention showing only the high pressure system.
  • FIG. 3 is an alternative construction for the engine of FIG. 1.
  • FIG. 4 shows in detail the joint between the shaft and tubular member of the engine.
  • FIG. 1 there is illustrated in a diagrammatic way, a by-pass gas turbine engine which includes compressor means 1, combustion equipment 2, turbine means 3, and a propulsion nozzle 4, all in flow series.
  • the engine illustrated is a three shaft engine in which the compressor means comprises low, intermediate, and high pressure compressors driven respectively by low, intermediate, and high pressure turbines each mounted on separate shafts. Since the invention has been applied only to the high pressure system of the engine, only this system is illustrated in detail. This is not meant to be restrictive on the scope of the invention, however, because the invention could be applied to either the intermediate or the low pressure systems, but less is to be gained in performance where the parts of the engine are not subject to the high differential thermal expansions, and high pressures, of the high pressure system.
  • Part of the air compressed by the low pressure compressor passes into a by-pass duct 5, by-passing the intermediate pressure and high pressure systems, and is mixed with the efflux from the low pressure turbine before passing to atmosphere through the propulsion nozzle 4.
  • FIG. 2 there is shown the high pressure system of a gas turbine engine to which the invention particularly relates.
  • the high pressure compressor 6, comprising six rotor stages 7 and six stator stages 8, supplies high pressure air via a diffuser 9, to the combustion equipment 2.
  • Fuel, supplied via burners 10, is burned in the combustion chamber 11 and the hot gases produced by combustion pass through a ring of nozzle guide vanes 12, to the rotor of the high pressure turbine 13. From the high pressure turbine 13 the gases pass to the remaining turbine stages of the engine which are not illustrated.
  • the high pressure turbine and compressor are drivingly interconnected by a shaft 14 which is supported at its ends in roller bearings 15 and 16 respectively.
  • the roller bearings allow for axial growth of the turbine, compressor and shaft assembly due to thermal expansion.
  • the turbine rotor 13 carries axially extending sealing ribs 17 which seal against radial faces of sealing members 19 attached to the nozzle guide vanes 12, which are supported from the engine casing 18.
  • the turbine and compressor assembly include a tubular member 20 which is located at its upstream end in a ball bearing 21 at the plane XX, thus providing a common datum from which expansion of the tubular member and the engine casing can be measured.
  • the member 20 is attached to the shaft 14 at a position between the compressor and the turbine by a coupling member 23 and serves to locate the shaft relative to plane XX.
  • the turbine expands rearwards, and the compressor forwards, relative to the coupling.
  • the coupling member 23, and its method of connection to the shaft and the tubular member 20, is shown in more detail in FIG. 4.
  • the member 20 is provided at its end with an external buttress thread 50 and with axially extending dogs 51.
  • the thread 50 engages with an internal thread on the coupling member 23, and the dogs 51 engage with corresponding dogs on a locking sleeve 52.
  • the sleeve 52 has radially extending splines 53 on its radially outer surface which engage with corresponding internal splines on the coupling member 23. The whole joint is held tight by a nut 54, which engages with further internal threads 55 on the coupling member 20.
  • the coupling member 20 is made from the same material as the shaft, or a material with a similar coefficient of expansion, so that the joint between the coupling member and the shaft can be a conventional curvic coupling.
  • An alternative method of joining the low expansion tubular member to the relatively high expansion shaft is to weld the coupling member 23 directly onto the end of the tubular member 20. This is preferably done by friction welding to avoid difficulties which can arise in conventional welding due to the different properties of the two materials.
  • a further alternative would be to friction weld a piece of the same material as the coupling member 23 onto the end of the member 20. This piece would have an external buttress thread 50 as above, and in this case, differential thermal expansion across the thread would be eliminated while retaining the adjusting feature.
  • the position of the joint between tubular member and the shaft can be calculated to give the desired sealing clearance at any one design condition of the engine provided that the temperatures of the shaft 14, tubular member 20, and casing 18 can be assessed to a reasonable degree of accuracy.
  • air is bled from a source in the compressor, the temperature of which is accurately known, and is fed through apertures 30 in the shaft 14 into an annular space 31 between the tubular member 20 and an additional sealing tube 32, to wash the outer surface of the tubular member.
  • the internal surface of the tubular member 20 is washed by a mixture of cooling air and oil, supplied through an oil tube 34, which has been used to cool the rear bearings, and the temperature of which can also be closely predicted.
  • the temperature of the casing 18 can be predicted quite accurately since the temperatures of the by-pass air and combustion chamber cooling air on its opposite sides do not vary widely.
  • the tubular member 20 is made from a material sold under the trade name of E.P.C. 10 by Henry Wiggin & Co. Ltd. and which is basically a Nickel Cobalt Iron alloy with a coefficient of expansion of between 0.000004 and 0.000005 per degree Centigrade.
  • the sealing clearance between the sealing ribs 17 and the nozzle guide vane can be maintained in the range 0.010 to 0.020 ins. at the design condition of the engine.
  • tubular member is present as described in FIG. 2, but in addition, the portion of the outer casing which surrounds the combustion chamber is made double-walled, and the outer wall is made from the low expansion material E.P.C. 10.
  • the outer wall 40 is made relatively thin, since it takes little loading, and expands rearwardly from the common datum XX.
  • the inner wall 41 is the equivalent of the casing wall 18 surrounding the combustion chamber in FIG. 2. It is made strong enough to support the compressor 6 and to contain the pressure inside the engine. It is anchored from a flange 42 at its downstream end so that it will expand axially in an upstream direction and it supports the compressor 6 by a sliding joint 43 at its upstream end.
  • the annular space 44 between the walls 40 and 41 is supplied with air bled from the compressor, the temperature and flow rate of which are known fairly accurately.
  • the air is exhausted into the by-pass passage 5 through apertures 45 in the inner wall 41 and apertures 46 in the engine casing downstream of the flange. Alternatively this air may be used for cooling of hot components of the turbine further downstream.
  • the use of the low expansion material in the outer wall reduces the amount of expansion of the casing and since the temperature and quantity of the air flow on both sides of the wall 40 are known, the temperature of the wall and hence its expansion can be calculated more easily.
  • the sealing clearance can be maintained in the range 0.005 ins. to 0.015 ins.
  • the whole or any part of the casing between datum XX and flange 42 may be made from a material of lower coefficient of thermal expansion than the inner wall. Due to the predictably low temperature of the outer wall which is washed on both sides by relatively cool air at known temperatures, it may not be necessary to use the very low expansion material E.P.C. for the outer wall. It may be possible, therefore, to use more titanium, which is lighter, and thus save some of the weight penalty incurred by the use of the doublewalled casing. This will depend on the expansion of the tubular member 20 and the temperature of the outer wall. Clearly, the use of the double wall for cooling the outer wall, in combination with the tubular member 20 is E.P.C. 1O material allows greater choice of materials for the outer wall, and other combinations of materials may be found which will give the required strength and sealing clearance.
  • the apertures in the casing through which the burners 10 pass must be sealed in a manner which allows for relative thermal expansion between the inner and outer casings.
  • a sleeve 50 screws into a threaded aperture in the inner casing, and a cup 51 is disposed between a shoulder 52 on the sleeve, and the inner casing.
  • the cup can slide axially relative to the sleeve, and radially relatively to the outer casing.
  • a gas turbine engine having a rotor member, a shaft, means for connecting the shaft and the rotor member, bearings for supporting the shaft with freedom for axial thermal expansion of the shaft, means for supporting a sealing member in static structure at a first radial plane of the engine, the rotor member being spaced from the sealing member by an axial sealing clearance, means for locating the shaft in static structure at a radial datum plane of the engine, said means comprising a member made of a material having a lower coefficient of thermal expansion than the shaft, means for connecting said member to the shaft and means for locating said member in static structure of the engine at said radial datum plane.
  • a gas turbine engine according to claim 1 and m which means are provided for bleeding air from a compressor of the engine and for passing said air over at least one surface of the member to control the temperature thereof.
  • a gas turbine engine according to claim 1 and in which the static structure comprises a casing of the engine and at least a portion of said casing between the first and second radial planes is made of a material having a lower coefficient of thermal expansion than the remainder of the casing between said planes.
  • a gas turbine engine in which at least a portion of the casing comprises an inner wall and an outer wall radially spaced therefrom, the outer wall supporting the sealing member and being made from said material, means being provided for supplying cooling air bled from a compressor of the engine to the space between the two walls.

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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)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US00169994A 1970-08-14 1971-08-09 Gas turbine engines Expired - Lifetime US3709637A (en)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
GB3917370 1970-08-14

Publications (1)

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US3709637A true US3709637A (en) 1973-01-09

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US00169994A Expired - Lifetime US3709637A (en) 1970-08-14 1971-08-09 Gas turbine engines

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US (1) US3709637A (enExample)
JP (1) JPS5217180B1 (enExample)
DE (1) DE2140337C3 (enExample)
FR (1) FR2102268B1 (enExample)
GB (1) GB1316452A (enExample)

Cited By (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3908361A (en) * 1972-12-16 1975-09-30 Rolls Royce 1971 Ltd Seal between relatively moving components of a fluid flow machine
US4578942A (en) * 1983-05-02 1986-04-01 Mtu Motoren-Und Turbinen-Union Muenchen Gmbh Gas turbine engine having a minimal blade tip clearance
US5165850A (en) * 1991-07-15 1992-11-24 General Electric Company Compressor discharge flowpath
US6053697A (en) * 1998-06-26 2000-04-25 General Electric Company Trilobe mounting with anti-rotation apparatus for an air duct in a gas turbine rotor
US20040126227A1 (en) * 2002-12-26 2004-07-01 Addis Mark E. Seal
US20120051886A1 (en) * 2010-08-30 2012-03-01 Leonard Paul Palmisano Locked spacer for a gas turbine engine shaft
RU2534680C1 (ru) * 2013-11-25 2014-12-10 Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации (Минпромторг России) Ротор турбомашины
US20170067360A1 (en) * 2015-09-09 2017-03-09 General Electric Technology Gmbh Steam turbine stage measurement system and a method

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10100642B2 (en) * 2015-08-31 2018-10-16 Rolls-Royce Corporation Low diameter turbine rotor clamping arrangement

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2622789A (en) * 1948-06-08 1952-12-23 Curtiss Wright Corp Turbine expansion section construction
US2962256A (en) * 1956-03-28 1960-11-29 Napier & Son Ltd Turbine blade shroud rings
US2992809A (en) * 1960-05-20 1961-07-18 Allis Chalmers Mfg Co Tandem compound steam turbine
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3514112A (en) * 1968-06-05 1970-05-26 United Aircraft Corp Reduced clearance seal construction
GB1238405A (enExample) * 1969-06-04 1971-07-07

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE281C (de) * 1877-07-03 E. VOSSKÖHLER, Werkstätten-Vorsteher der Bergisch-Märkischen Eisenbahn in Düsseldorf Vorrichtung für Nothsignale an Eisenbahnwagen
GB586200A (en) * 1944-07-21 1947-03-11 Karl Baumann Improvements in bladed drum-type rotor constructions operating under high temperature conditions
FR1404212A (fr) * 1964-08-12 1965-06-25 Bbc Brown Boveri & Cie Machine à ondes de choc

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2622789A (en) * 1948-06-08 1952-12-23 Curtiss Wright Corp Turbine expansion section construction
US2962256A (en) * 1956-03-28 1960-11-29 Napier & Son Ltd Turbine blade shroud rings
US2992809A (en) * 1960-05-20 1961-07-18 Allis Chalmers Mfg Co Tandem compound steam turbine
US3391904A (en) * 1966-11-02 1968-07-09 United Aircraft Corp Optimum response tip seal
US3514112A (en) * 1968-06-05 1970-05-26 United Aircraft Corp Reduced clearance seal construction
GB1238405A (enExample) * 1969-06-04 1971-07-07

Cited By (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3908361A (en) * 1972-12-16 1975-09-30 Rolls Royce 1971 Ltd Seal between relatively moving components of a fluid flow machine
US4578942A (en) * 1983-05-02 1986-04-01 Mtu Motoren-Und Turbinen-Union Muenchen Gmbh Gas turbine engine having a minimal blade tip clearance
US5165850A (en) * 1991-07-15 1992-11-24 General Electric Company Compressor discharge flowpath
US6053697A (en) * 1998-06-26 2000-04-25 General Electric Company Trilobe mounting with anti-rotation apparatus for an air duct in a gas turbine rotor
US20040126227A1 (en) * 2002-12-26 2004-07-01 Addis Mark E. Seal
US6910858B2 (en) * 2002-12-26 2005-06-28 United Technologies Corporation Seal
US20120051886A1 (en) * 2010-08-30 2012-03-01 Leonard Paul Palmisano Locked spacer for a gas turbine engine shaft
US8967977B2 (en) * 2010-08-30 2015-03-03 United Technologies Corporation Locked spacer for a gas turbine engine shaft
RU2534680C1 (ru) * 2013-11-25 2014-12-10 Российская Федерация, от имени которой выступает Министерство промышленности и торговли Российской Федерации (Минпромторг России) Ротор турбомашины
US20170067360A1 (en) * 2015-09-09 2017-03-09 General Electric Technology Gmbh Steam turbine stage measurement system and a method
US10267178B2 (en) * 2015-09-09 2019-04-23 General Electric Company Steam turbine stage measurement system and a method

Also Published As

Publication number Publication date
GB1316452A (en) 1973-05-09
FR2102268B1 (enExample) 1974-03-29
JPS5217180B1 (enExample) 1977-05-13
FR2102268A1 (enExample) 1972-04-07
DE2140337C3 (de) 1975-11-20
DE2140337B2 (de) 1975-04-10
DE2140337A1 (de) 1973-07-26

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