EP1717415B1 - Turbinenmodul für einen Gasturbinentriebwerk - Google Patents

Turbinenmodul für einen Gasturbinentriebwerk Download PDF

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Publication number
EP1717415B1
EP1717415B1 EP06113242A EP06113242A EP1717415B1 EP 1717415 B1 EP1717415 B1 EP 1717415B1 EP 06113242 A EP06113242 A EP 06113242A EP 06113242 A EP06113242 A EP 06113242A EP 1717415 B1 EP1717415 B1 EP 1717415B1
Authority
EP
European Patent Office
Prior art keywords
module
compressor
bolts
turbine
flange
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.)
Active
Application number
EP06113242A
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English (en)
French (fr)
Other versions
EP1717415A1 (de
Inventor
Michel Brault
Maurice Judet
Patrick Pasquis
Thomas Langevin
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.)
Safran Aircraft Engines SAS
Original Assignee
SNECMA SAS
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 SNECMA SAS filed Critical SNECMA SAS
Publication of EP1717415A1 publication Critical patent/EP1717415A1/de
Application granted granted Critical
Publication of EP1717415B1 publication Critical patent/EP1717415B1/de
Active 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
    • 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
    • F01D5/027—Arrangements for balancing
    • 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
    • F01D5/06—Rotors for more than one axial stage, e.g. of drum or multiple disc type; Details thereof, e.g. shafts, shaft connections
    • F01D5/066—Connecting means for joining rotor-discs or rotor-elements together, e.g. by a central bolt, by clamps
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05B—INDEXING SCHEME RELATING TO WIND, SPRING, WEIGHT, INERTIA OR LIKE MOTORS, TO MACHINES OR ENGINES FOR LIQUIDS COVERED BY SUBCLASSES F03B, F03D AND F03G
    • F05B2260/00—Function
    • F05B2260/30—Retaining components in desired mutual position
    • F05B2260/301—Retaining bolts or nuts
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • 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
    • F05D2260/00—Function
    • F05D2260/30—Retaining components in desired mutual position
    • F05D2260/34—Balancing of radial or axial forces on regenerative rotors

Definitions

  • the present invention relates to the field of gas turbine engines and aims in particular a turbine module, in particular the high pressure module.
  • An axial gas turbine engine includes a rotary compressor assembly, supplying a stationary combustion chamber which itself delivers hot gases to a rotating turbine assembly.
  • a turbine rotor integral with a compressor rotor forms a body and the motor may comprise one or more bodies, generally two or three rotating at different speeds.
  • a dual-body engine comprises a low pressure body LP and an HP high pressure body.
  • its architecture is often divided into modules. For example, for the high-pressure body of a double-body engine, all the parts constituting the compressor are arranged in an HP compressor module and all the parts constituting the turbine rotor into an HP turbine module.
  • This latter module consists of only rotating parts, for example a turbine disc on which are bolted a labyrinth sealing assembly and a ferrule with upstream assembly flange, and a sealing assembly associated with a downstream bearing. .
  • the HP compressor and HP turbine modules are assembled via specific flanges; these flanges transmit the engine torque of the turbine to the compressor.
  • the connection by these flanges must be sufficiently resistant to fulfill this function.
  • US 5,816,776 for example, describes this state of the art.
  • the modules must be provided to the assembly being perfectly balanced in rotation.
  • an upstream balancing plan and a downstream balancing plan are provided.
  • the balancing plane is that in which balancing weights are available at a distance from the axis and at a given angle.
  • the flange that is at the The boundary of the module constitutes an appropriate balancing plan. Balancing weights are therefore available in the area of the boundary flange. This is particularly the case at the boundary between the HP compressor module and the HP turbine module.
  • Each module is presented to the assembly being balanced in this way.
  • the applicant has therefore set itself the goal of developing a turbine module that satisfies these constraints.
  • the gas turbine engine turbine module comprising at least one turbine disk and a disk-shaped component mounted upstream on the turbine disk with respect to the gas flow, said module comprising a means of assembly between the compressor of said engine and the turbine disk, is characterized in that the component, before assembly of the module to the compressor, is pre-assembled to the turbine disk, by bolting to a fastening flange integral with the turbine disk.
  • the means for assembling the compressor to the turbine disk comprises said fixing flange.
  • the component is disc-shaped with a web between its hub and its periphery, and is mounted to said flange by bolting through first holes in the web, second holes are formed in the web for assembly by bolting from the turbine module to the compressor.
  • the component is fixed to said flange by bolts whose number is between 2 and 8 distributed over the circumference.
  • the bolts are retained by nuts crimped on the flange, downstream side, together with nuts for assembly of the module to the compressor.
  • the fastening bolts of the component more particularly have a specific shaped head which differs from the heads of the fastening bolts of the compressor.
  • the component comprises an upstream balancing device disposed on said component independently of the assembly means to the compressor.
  • the invention applies in particular to a module whose component is a support disc of labyrinth sealing elements.
  • the invention also covers the compressor associated with the module to form a gas turbine engine, comprising a downstream flange for attachment to the module. On this flange housing or passages are formed for the heads of bolts of fixings already in place on the module.
  • This module comprises a turbine disk 3, with a hub with axial section increasing in thickness close to the axis, and at the periphery of which are mounted the turbine blades 4. These are housed in axial recesses formed on the rim of the disc.
  • a component 5 is mounted on the disc 3 upstream, that is to say on the left in the figure. Upstream and downstream are defined with respect to the flow of gases in the engine.
  • the component is here a symmetrical disk of revolution relative to the axis of the machine.
  • This disc comprises a hub of increasing thickness towards the axis of the motor, a portion forming a web 51 towards the periphery.
  • the disc at its periphery carries radial annular blades 53 forming the rotating portion of labyrinth seals. Their counterpart is not represented.
  • the disc 5 is fixed to the disc 3 by bolting to a fastening flange 31 secured to the disc upstream thereof.
  • the bolts 7 comprise a head 71, a rod 72 passing through an orifice formed in the web 51 and a bore machined in the flange 31, and cooperate with a nut 73.
  • a disc 6 On the downstream side of the module there is a disc 6 with a journal 61 forming a support for a downstream bearing 62.
  • the disk is bolted to a downstream mounting flange 33.
  • the bolts 64 are fixed all around the disc. It also has sealing elements 63 for a labyrinth seal.
  • Disk 6 forms a downstream balancing plane. Balancing weights are fitted with the fixing bolts.
  • the module as represented on the figure 1 is pre-assembled ready to be assembled and assembled to a compressor.
  • the function of the bolts 7 is to maintain the component 5 integral with the turbine disk during handling.
  • the bolts 7 are preferably four in number and equidistant on the circumference. Their number can be between 2 and 8. In fact, you have to take into account the fastening bolts of the compressor to the disc. The bolts 7 do not come to disturb the tiering between the disc 3 and the disc 5.
  • FIG. 2 the bolting detail is seen after the turbine module has been attached to a compressor attachment flange 9.
  • the latter is not represented.
  • the fixing flange 9 is annular and constitutes the downstream boundary of the compressor.
  • the figure 2 represents a partial axial section taken along a bolt 7 for fixing the component 5 to the disc 3.
  • the figure 3 represents a partial axial section taken along a bolt 8 for fixing the flange 9 to the disk 3.
  • the heads 71 of the bolts 7 are engaged in wide scallops of the flange 9 in such a way that they engage the disc of the component 5 directly. These bolts therefore have no effect on the connection of the flange 9.
  • Bolts 8 of the figure 3 each participate in fixing the flange 9 to the disc 3.
  • the heads 81 bear against the upstream surface of the flange 9 which is held tight against the disk 5 by the nut 83 applied against the downstream face of the flange 31.
  • the rod 82 of the bolt 8 passes through the disk of the component 5 and the two flanges 9 and 31.
  • annular flange 9 is shown as seen from the upstream side. It comprises notches in the form of scallops to allow it to bypass the heads 71 of the bolts 7, and abut against the disk of the component 5. It also includes holes for the passage of the rods 82 of the bolts 8, whose head 81 comes to bear on the edge of the holes.
  • the component 5 has the same holes in the sail 51 capable of passing either the bolts 7 or the bolts 8.
  • this type of head while also providing the keying, has the further advantage of avoiding seizing, at the time of the modular disassembly of the compressor and the turbine, the bolts 8. If the bolts 8 remain blocked, the specific heads 81 of the bolts 8 are capable of withstanding a violent over-torque for shearing the shaft 82 of the head 81. In all cases, the modular disassembly is ensured.
  • the flange 31 is shown from the downstream side.
  • the nuts 73 and 83 are preferably mounted and crimped on site before assembling the parts. It is noted that they have a tongue 731 respectively 831 of rectangular shape so as to ensure mutual locking in rotation. This facilitates the screwing of bolts 7 and 8 during assembly.
  • the nuts 73 and 83 crimped on the flange 31 are identical.
  • fastening bolts 8 and 7 are in a concrete example of 28 and 4, respectively. It was verified that the 28 bolts could safely secure the attachment of the compressor to the disc. The difference of 4 with respect to an assembly at 32 does not affect. Calculations have demonstrated this in particular for: the passage torque, mechanical stresses, the life of parts, etc ....
  • a module comprising this balancing device is represented on the figure 7 .
  • the module comprises a flange 10 for hooking the weights on this face.
  • the flange 10 is annular with a face perpendicular to the axis of the module, and comprises a plurality of machining for assembling weights.
  • the number of machining is preferably equal to that of the number of fins mounted on the turbine disk.
  • the figure 7 shows a weight 11 in place held by bolts 111. In this case, the installation of these weights allows the return of the upstream balancing plane rotor HP turbine module. This contributes to one of the delivery conditions of a so-called clean module.
  • a module according to the present invention does not degrade the maintenance of the surrounding modules.

Landscapes

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

Claims (12)

  1. Turbinenmodul für einen Gasturbinenmotor, das wenigstens eine Turbinenscheibe (3) und ein scheibenförmiges Bauteil (5) aufweist, das an der Turbinenscheibe bezüglich der Gasabströmung stromaufwärts angebracht ist, wobei das Modul ein Mittel zum Zusammenfügen des Motors und des Kompressors aufweist, dadurch gekennzeichnet, daß das Bauteil (5) vor dem Zusammenfügen des Moduls und des Kompressors durch eine Bolzenverbindung an einem Befestigungsflansch (31) vorher zusammengefügt ist, der mit der Turbinenscheibe (3) fest verbunden ist.
  2. Modul nach Anspruch 1, bei welchem das Mittel zum Zusammenfügen an dem Kompressor den Befestigungsflansch (31) aufweist.
  3. Modul nach Anspruch 2, bei welchem das Bauteil scheibenförmig ist mit einer Abdeckung (51) zwischen seiner Nabe und seinem Umfang und an dem Flansch (31) durch eine Bolzenverbindung (7) durch erste Bohrungen in der Abdeckung (51) angebracht ist, wobei die zweiten Bohrungen in der Abdeckung zum Zusammenfügen des Turbinenmoduls (3) und des Kompressors durch eine Bolzenverbindung (8) angeordnet sind.
  4. Modul nach Anspruch 3, bei welchem das Bauteil (5) die gleichen Bohrungen in der Abdeckung (51) besitzt, die in der Lage ist, entweder die Bolzen (7) oder die Bolzen (8) durchzulassen.
  5. Modul nach Anspruch 3 oder 4, bei welchem das Bauteil an dem Flansch (31) durch Bolzen (7) befestigt ist, deren Anzahl zwischen 2 und 8 liegt und die am Umfang verteilt sind.
  6. Modul nach einem der Ansprüche 2 bis 5, bei welchem die Bolzen (7) durch Muttern (73) gehalten werden, die an dem Flansch (31) an der stromabwärts gelegenen Seite zusammen mit Muttern (83) zum Zusammenfügen des Moduls und dem Kompressor gehalten werden.
  7. Modul nach Anspruch 6, bei welchem die an dem Flansch (31) eingefaßten Muttern (73) und (83) identisch sind.
  8. Modul nach einem der Ansprüche 2 bis 5, bei welchem die Bolzen (7) zur Befestigung des Bauteils einen spezifischen Kopf haben, der sich von den Köpfen der Bolzen (8) zur Befestigung des Kompressors unterscheidet.
  9. Modul nach einem der Ansprüche 2 bis 8, das eine stromaufwärts gelegene Ausgleichsvorrichtung (10, 11) aufweist, die an dem Bauteil (5) unabhängig von dem Mittel zum Zusammenfügen und des Kompressors angeordnet ist.
  10. Modul nach einem der Ansprüche 2 bis 9, bei welchem das Bauteil (5) eine Trägerscheibe für Labyrinthdichtungselemente ist.
  11. Einheit mit einem Kompressor und einem Modul nach einem der vorhergehenden Ansprüche, um einen Gasturbinenmotor zu bilden, wobei der Kompressor einen stromabwärts gelegenen Flansch (9) zur Befestigung an dem Modul aufweist, an welchem Aufnahmen oder Durchgänge für die Befestigungsbolzenköpfe (71) vorgesehen sind, die sich bereits an Platz an dem Modul befinden.
  12. Gasturbinenmotor mit einem Turbinenmodul nach einem der Ansprüche 1 bis 10.
EP06113242A 2005-04-29 2006-04-27 Turbinenmodul für einen Gasturbinentriebwerk Active EP1717415B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR0551124A FR2885167B1 (fr) 2005-04-29 2005-04-29 Module de turbine pour moteur a turbine a gaz

Publications (2)

Publication Number Publication Date
EP1717415A1 EP1717415A1 (de) 2006-11-02
EP1717415B1 true EP1717415B1 (de) 2008-06-11

Family

ID=35169522

Family Applications (1)

Application Number Title Priority Date Filing Date
EP06113242A Active EP1717415B1 (de) 2005-04-29 2006-04-27 Turbinenmodul für einen Gasturbinentriebwerk

Country Status (9)

Country Link
US (1) US7364402B2 (de)
EP (1) EP1717415B1 (de)
CN (1) CN1854485B (de)
CA (1) CA2544784C (de)
DE (1) DE602006001430D1 (de)
ES (1) ES2310887T3 (de)
FR (1) FR2885167B1 (de)
RU (1) RU2403401C2 (de)
UA (1) UA88281C2 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11092012B2 (en) 2017-03-27 2021-08-17 MTU Aero Engines AG Turbomachine component arrangement

Families Citing this family (31)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8167547B2 (en) * 2007-03-05 2012-05-01 United Technologies Corporation Gas turbine engine with canted pocket and canted knife edge seal
FR2914008B1 (fr) * 2007-03-21 2009-10-09 Snecma Sa Ensemble rotatif d'une soufflante de turbomachine
FR2921422B1 (fr) * 2007-09-26 2009-12-18 Snecma Piece annulaire de turbomachine portant des ecrous a jupe de sertissage
FR2931869B1 (fr) * 2008-05-29 2014-12-12 Snecma Bride annulaire de fixation d'un element de rotor ou de stator
FR2931873B1 (fr) * 2008-05-29 2010-08-20 Snecma Ensemble d'un disque de turbine d'un moteur a turbine a gaz et d'un tourillon support de palier,circuit de refroidissement d'un disque de turbine d'un tel ensemble.
FR2938292B1 (fr) * 2008-11-07 2010-12-24 Snecma Bride annulaire de fixation d'un element de rotor ou de stator dans une turbomachine
FR2939836B1 (fr) * 2008-12-12 2015-05-15 Snecma Joint d'etancheite de plateforme dans un rotor de turbomachine
FR2946083B1 (fr) * 2009-05-28 2011-06-17 Snecma Turbine basse-pression
FR2974865B1 (fr) * 2011-05-04 2013-07-05 Snecma Rotor haute pression pour turbomachine d'aeronef, comprenant des moyens de detrompage associes a des boulons de prefixation de module de turbine
BR112015007733B1 (pt) 2012-10-08 2022-05-03 United Technologies Corporation Motores de turbina a gás, e, método para distribuir peso entre um conjunto de propulsor e um conjunto de gerador de gás de um motor de turbina a gás
FR3001515B1 (fr) * 2013-01-25 2015-03-20 Snecma Assemblage de masselotte d'equilibrage a un element de rotor
FR3002585B1 (fr) * 2013-02-27 2016-07-15 Snecma Equilibrage d'un ensemble rotatif dans une turbomachine
EP2986824B1 (de) 2013-04-18 2020-05-27 United Technologies Corporation Turbinenminischeibenstossfänger für einen gasturbinenmotor
EP2824278B1 (de) * 2013-07-08 2018-11-07 MTU Aero Engines GmbH Vorrichtung, Schaufel-Vorrichtungs-Verbund, Verfahren sowie Strömungsmaschine
DE102013216377A1 (de) 2013-08-19 2015-03-12 Rolls-Royce Deutschland Ltd & Co Kg Verfahren zum Auswuchten und zur Montage eines Turbinenrotors
RU2532390C1 (ru) * 2013-09-10 2014-11-10 Открытое акционерное общество "Авиадвигатель" Ротор турбины высокого давления
FR3021066B1 (fr) * 2014-05-19 2019-05-10 Safran Aircraft Engines Disque de rotor equilibre, et procede d'equilibrage
ES2711912T3 (es) 2015-04-16 2019-05-08 Thilo Kraemer Dispositivo de ensayo para la ejecución de mediciones de dureza
DE202015101878U1 (de) 2015-04-16 2015-04-30 Thilo Kraemer Prüfgerät
EP3091179B1 (de) * 2015-05-07 2021-06-30 MTU Aero Engines AG Rotoranordnung für eine strömungsmaschine und verdichter
US11143467B2 (en) 2015-05-20 2021-10-12 Other Lab, Llc Membrane heat exchanger system and method
CN105401981A (zh) * 2015-12-29 2016-03-16 中国航空工业集团公司沈阳发动机设计研究所 一种高转速低压涡轮转子结构
CN107932060B (zh) * 2017-11-23 2019-03-22 中国航发沈阳黎明航空发动机有限责任公司 一种控制燃机低压一级涡轮盘磨损的方法
FR3078363B1 (fr) * 2018-02-23 2021-02-26 Safran Aircraft Engines Anneau mobile d'etancheite
US11448081B2 (en) 2019-10-18 2022-09-20 Raytheon Technologies Corporation Balanced circumferential seal
CN111485955A (zh) * 2020-04-16 2020-08-04 中国航发沈阳发动机研究所 一种转子轮盘装配结构
US11549373B2 (en) 2020-12-16 2023-01-10 Raytheon Technologies Corporation Reduced deflection turbine rotor
US11578599B2 (en) * 2021-02-02 2023-02-14 Pratt & Whitney Canada Corp. Rotor balance assembly
FR3119647B1 (fr) * 2021-02-11 2023-01-13 Safran Aircraft Engines Procédé de réparation d’une bride de rotor de turbomachine
RU208145U1 (ru) * 2021-06-07 2021-12-06 Публичное Акционерное Общество "Одк-Сатурн" Узел ротора турбины высокого давления
CN114215611B (zh) * 2021-12-01 2023-07-14 东方电气集团东方汽轮机有限公司 一种燃气轮机透平动叶轴向定位用气封装配体

Family Cites Families (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5275534A (en) * 1991-10-30 1994-01-04 General Electric Company Turbine disk forward seal assembly
US5562404A (en) * 1994-12-23 1996-10-08 United Technologies Corporation Vaned passage hub treatment for cantilever stator vanes
FR2744761B1 (fr) * 1996-02-08 1998-03-13 Snecma Disque labyrinthe avec raidisseur incorpore pour rotor de turbomachine
US6893222B2 (en) * 2003-02-10 2005-05-17 United Technologies Corporation Turbine balancing
FR2857419B1 (fr) * 2003-07-11 2005-09-23 Snecma Moteurs Liaison amelioree entre disques aubages sur la ligne rotor d'un compresseur

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11092012B2 (en) 2017-03-27 2021-08-17 MTU Aero Engines AG Turbomachine component arrangement

Also Published As

Publication number Publication date
US7364402B2 (en) 2008-04-29
UA88281C2 (ru) 2009-10-12
DE602006001430D1 (de) 2008-07-24
RU2403401C2 (ru) 2010-11-10
US20070059164A1 (en) 2007-03-15
RU2006114654A (ru) 2007-11-10
CA2544784C (fr) 2014-01-28
CA2544784A1 (fr) 2006-10-29
CN1854485A (zh) 2006-11-01
ES2310887T3 (es) 2009-01-16
FR2885167B1 (fr) 2007-06-29
CN1854485B (zh) 2011-05-25
FR2885167A1 (fr) 2006-11-03
EP1717415A1 (de) 2006-11-02

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