EP0192556A1 - Turbinengehäuse mit einer Einrichtung, um den Abstand zwischen den Schaufelspitzen und dem Gehäuse einzustellen - Google Patents

Turbinengehäuse mit einer Einrichtung, um den Abstand zwischen den Schaufelspitzen und dem Gehäuse einzustellen Download PDF

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Publication number
EP0192556A1
EP0192556A1 EP86400286A EP86400286A EP0192556A1 EP 0192556 A1 EP0192556 A1 EP 0192556A1 EP 86400286 A EP86400286 A EP 86400286A EP 86400286 A EP86400286 A EP 86400286A EP 0192556 A1 EP0192556 A1 EP 0192556A1
Authority
EP
European Patent Office
Prior art keywords
wall
casing
segments
turbomachine
clearance
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.)
Granted
Application number
EP86400286A
Other languages
English (en)
French (fr)
Other versions
EP0192556B1 (de
Inventor
Jean-Paul Lagrange
Jean-Max Marie Silhouette
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
Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
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 Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA, SNECMA SAS filed Critical Societe Nationale dEtude et de Construction de Moteurs dAviation SNECMA
Publication of EP0192556A1 publication Critical patent/EP0192556A1/de
Application granted granted Critical
Publication of EP0192556B1 publication Critical patent/EP0192556B1/de
Expired 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
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • 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
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/16—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means
    • F01D11/18—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing by self-adjusting means using stator or rotor components with predetermined thermal response, e.g. selective insulation, thermal inertia, differential expansion

Definitions

  • the invention relates to a turbomachine casing associated with a device making it possible to adjust the clearance between said casing and the movable blades of a rotor in operation, automatically.
  • a turbomachine casing associated with a device for adjusting the clearance between movable blades and casing according to the invention is characterized in that said casing is internally lined with a rigid annular wall connected to the casing by connecting means which allow said wall completely free to move in the radial direction according to the expansions / contractions in operation and in that said wall consists of a succession of integral segments, each segment being alternately of a type with low thermal inertia and of a type with high thermal inertia so that the radial displacements of the said wall adapt in all operating conditions of the turbomachine to the radial displacements of the moving blade heads.
  • the inner wall segments with low thermal inertia are thin, their internal surface is directly in contact with the gases of the vein. circulating through the stage of movable blades considered, their external and lateral surfaces are coated with a thermally insulating material and the adjacent segments of inner wall with high thermal inertia are very thick and coated on all their surfaces with material thermally insulating.
  • the temperature T of the hot engine gases which pass through a grid of blades such as 1 creates a flow of heat which, passing through the blades, spreads as far as the disc which carries them and causes a radial expansion of the assembly, therefore a displacement of the blade heads 1.
  • the radius of gyration R of the blade heads follows a law which can be represented by the relation: binding R to R o , gyratory radius of the blade heads at T o , ambient temperature on the ground, K 1 being the coefficient of radial thermal expansion of the rotor.
  • each grid of movable blades such as 1 can however be characterized fairly precisely by means of four parameters.
  • the speed of rotation N of the rotor creates a centrifugal force which acts on the whole of the rotor and causes another radial displacement of the blade heads 1.
  • T of the hot engine gases at a given point in the stream is a function of the speed of rotation N of the rotor: where K 3 is the coefficient of proportionality between T and N 2.
  • the radius of gyration R of the movable blades 1 is linked to the temperature of the hot..engine gases, at a given point in the vein, by a relationship of simultaneity:
  • the evolution of the temperature T of the engine gases is influenced by the fact that the evolution of the speed N is due to a temporary excess or deficit of the flow rate of burned fuel in the combustion chamber in relation to the flow required in stabilized operation.
  • ⁇ T c is the temperature difference due to the excess or deficit of fuel burned
  • T N is the temperature that the turbine gases would have if the speed N were stabilized.
  • the temperature difference ⁇ T c intervenes directly in the evolution of the radius of gyration R due to the temperature T as before.
  • its duration is at most equal to that of the speed transient N, ie 5 to 10 seconds maximum for a simple evolution. His influence is therefore noticeable only on the expansion of the movable blades 1, that is to say, in the case of a "unit step" of temperature as above, in connection only with the gain K'l and the time constant
  • FIG. 1 represents, in transverse section with respect to the axis of rotation of the turbomachine, a simplified view of an embodiment of the invention.
  • a turbomachine Facing the movable blade heads 1, there is a fixed part of a turbomachine constituted by a stator casing 2 which in the example shown is in two parts 2a, 2b, each having a generally semi-cylindrical shape.
  • Each part, respectively 2a and 2b, carries at its ends flanges, respectively 3a and 4a, 3b and 4b which are assembled by any known means, such as bolting.
  • the casing 2 is internally lined with a rigid wall 5, which in the example shown is also in two parts 5a and 5b.
  • This inner wall 5 consists of a succession of integral segments 6 which are of two different types 6a and 6b, more clearly shown in Figure 2, and arranged alternately.
  • a segment such as 6a is thin, its internal surface 6i is directly in contact with the gases of the vein circulating in the stage of movable blades 1 and its external and lateral surfaces are coated with a layer 7a of a material thermally insulating.
  • These segments 6a of the inner wall 5 therefore very quickly take the temperature of the gases of the vein.
  • An adjacent segment such as 6b is very thick and all of its surfaces are coated with a layer 7b of a thermally insulating material.
  • segments 6b of the inner wall 5 thus have a high thermal inertia and their thermal connection with the outside takes place almost only through their junctions with the adjacent segments 6a. They therefore take the temperature of the vein gases very slowly.
  • the insulating layers 7a and 7b are flexible enough to follow all the thermal expansions / contractions of the interior wall 5.
  • the segments 6a and pb are in sufficient number so that the initially circular shape of the annular wall is sufficiently well preserved during the expansions / thermal contractions occurring during the operation of the turbomachine.
  • the internal surface of the inner wall 5, both for the segments 6a and for the adjacent segments 6b, can be covered with a layer 8 of abradable material constituting a wear and seal lining susceptible during operation of come into contact with the ends of movable blades 1 without causing damage.
  • This material is determined so as not to create a thermal barrier between the internal surface 6i of the segments 6a of the internal wall 5 and the gases of the vein and not to slow down the thermal expansions / contractions of the internal wall 5.
  • Each part, respectively 5a and 5b, of the inner wall 5 is fixed inside the corresponding part of the casing 2, respectively 2a and 2b, by means of connecting rods 9.
  • a yoke 10 is fixed, for example by screwing, on at least some of the segments 6b of the inner wall 5, at a 7th end of these segments in a radially external zone.
  • a yoke 11 is also fixed, for example by screwing, on the internal surface of the casing 2, in a position circumferentially offset with respect to the associated yoke 10.
  • Each link 9 is provided at its ends with yokes, respectively 9a and 9b which cooperate by means of axes of rotation 10a and lla with said yokes 10 and 11.
  • links 9 are thus placed in a direction substantially tangential to the inner wall 5 and thus leave the wall 5 free to move radially under the influence of thermal expansion / contraction.
  • an access hole 5c can be made in the interior wall 5, preferably at the level of a thin segment 6a.
  • each end of the inner wall part, respectively 5a and 5b comprises a half-segment respectively 6c or 6d of the very thick type. As shown in more detail in FIGS. 3, 3a and 3b, these half-segments 6c and 6d are joined by their respective end faces by means, for example, of a bolt 12.
  • each of these faces comprises by example a tenon 13 and a mortise 14 arranged in two perpendicular directions and cooperating respectively with a mortise 14a and a tenon 13a of the face of the associated half-segment to very precisely join the two parts 5a and 5b of the inner wall 5.
  • a passage access 15 is provided through the casing 2 to allow the establishment of the bolting 12.
  • the internal wall 5 is placed in a housing constituted by an annular recess 16 formed on the internal face of the casing 2. Under the action of the pressure P of the gases, the internal wall 5 thus comes to be pressed laterally on the surface 16a of the recess 16 where the pressure P is the lowest.
  • the corresponding lateral surface of the wall 5 is coated with a layer 7b of thermally insulating material, as previously described, which in this zone prevents gas leaks, reduces contact friction and reduces heat exchange between the interior wall 5 and the casing 2.
  • a layer 7b of thermally insulating material as previously described, which in this zone prevents gas leaks, reduces contact friction and reduces heat exchange between the interior wall 5 and the casing 2.
  • the internal wall 5 is pressed against a surface 16a of the recess 16 located downstream with respect to the direction of gas flow in the stream of the turbomachine.
  • the inner wall 5 would be pressed against an upstream surface.
  • the inner wall 5 like the casing 2 have a generally cylindrical shape which corresponds to the outer shape of the gas stream of the turbomachine in the area considered. But of course, the invention applies in the same way in the case where this shape of vein is conical and in this case the inner wall 5 also has a generally conical shape adapted to the vein.
  • the solution proposed by the invention consists in producing a "thermal model" of the rotor on the stator.
  • an interior wall 5 is produced which, in regimes transient as well as in stabilized regimes, very precisely follows the radial movements of the blades of the rotor, and this by the only thermal effect on this inner wall 5 of the hot gases which lick it. Because this interior wall 5 is a complete circumference, any peripheral expansion results in a radial expansion of the interior wall 5. This is the principle used.
  • FIG. 5 Certain possible adaptations are shown in FIG. 5 for a segment 106b equivalent to a segment 6b.
  • segments 206b with high thermal inertia radially spaced apart towards the outside.
  • the adjacent segments 206a with low thermal inertia each extend respectively by a portion 25.
  • These portions 25 remain without influence on the changes in diameter of the interior wall 5 because they are separated by a slot 26 which can have different shapes , as shown in FIGS. 6a, right, 6b, oblique, or 6c, with a balonette.
  • the ends of the parts 25 are covered by a part 27 which covers the slot 26.
  • each element 105 of internal wall 5 is terminated by two half-segments 306c and 306d of the type with high thermal inertia to facilitate the attachment of the elements to each other.
  • the rigidity of the interior wall 5 may be insufficient despite its pressing on a lateral surface 16a of the recess 16 of the casing 2 (see FIG. 4).
  • FIG. 8a are placed on the upstream and downstream side edges of the segments 406a of the inner wall 5, two ribs 19 thin enough to maintain the thermal performance of the two types of segments.
  • the requested lateral seal can then be produced on one of these ribs, bearing on the lateral surface 16a of the recess 16 of the casing 2 (see FIG. 4).
  • FIG. 8b it is possible, as an alternative, to provide on each lateral edge of a thin segment 506a stiffening elements 20 and 21 attached and fixed to the external wall of the segments 506a.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
EP86400286A 1985-02-13 1986-02-11 Turbinengehäuse mit einer Einrichtung, um den Abstand zwischen den Schaufelspitzen und dem Gehäuse einzustellen Expired EP0192556B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8502022 1985-02-13
FR8502022A FR2577281B1 (fr) 1985-02-13 1985-02-13 Carter de turbomachine associe a un dispositif pour ajuster le jeu entre aubes mobiles et carter

Publications (2)

Publication Number Publication Date
EP0192556A1 true EP0192556A1 (de) 1986-08-27
EP0192556B1 EP0192556B1 (de) 1989-05-24

Family

ID=9316218

Family Applications (1)

Application Number Title Priority Date Filing Date
EP86400286A Expired EP0192556B1 (de) 1985-02-13 1986-02-11 Turbinengehäuse mit einer Einrichtung, um den Abstand zwischen den Schaufelspitzen und dem Gehäuse einzustellen

Country Status (5)

Country Link
US (1) US4787817A (de)
EP (1) EP0192556B1 (de)
JP (1) JPS61190101A (de)
DE (1) DE3663556D1 (de)
FR (1) FR2577281B1 (de)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2244524A (en) * 1990-05-31 1991-12-04 Gen Electric Clearance control in gas turbine engines
EP0716220A1 (de) * 1994-12-07 1996-06-12 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Ungeteilter Leitkranz für eine Turbomaschine
EP0770761A1 (de) * 1995-10-23 1997-05-02 United Technologies Corporation Mantelring zur Schaufelspitzenabdichtung

Families Citing this family (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5127795A (en) * 1990-05-31 1992-07-07 General Electric Company Stator having selectively applied thermal conductivity coating
US6240727B1 (en) * 2000-04-27 2001-06-05 The United States Of America As Represented By The Secretary Of The Navy Manufacture of Nitinol rings for thermally responsive control of casing latch
GB0218060D0 (en) * 2002-08-03 2002-09-11 Alstom Switzerland Ltd Sealing arrangements
DE10305899B4 (de) * 2003-02-13 2012-06-14 Alstom Technology Ltd. Dichtungsanordnung zur Dichtspaltreduzierung bei einer Strömungsrotationsmaschine
DE102007054483A1 (de) * 2007-11-15 2009-05-20 Mtu Aero Engines Gmbh Bauteil mit ringartiger oder rohrartiger Form
GB2462581B (en) * 2008-06-25 2010-11-24 Rolls Royce Plc Rotor path arrangements
WO2014143311A1 (en) * 2013-03-14 2014-09-18 Uskert Richard C Turbine shrouds
DE102013212741A1 (de) * 2013-06-28 2014-12-31 Siemens Aktiengesellschaft Gasturbine und Hitzeschild für eine Gasturbine
EP3375980B1 (de) * 2017-03-13 2019-12-11 MTU Aero Engines GmbH Dichtungsträger für eine strömungsmaschine
CN112855352B (zh) * 2019-11-28 2022-03-22 中国航发商用航空发动机有限责任公司 高压涡轮实时叶尖间隙的计算方法和控制方法

Citations (7)

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Publication number Priority date Publication date Assignee Title
FR1219504A (fr) * 1958-03-25 1960-05-18 Zd Y V I Bague de joint étanche pour roue à aubes de turbines à gaz
CH408960A (de) * 1961-09-04 1966-03-15 Licentia Gmbh Radialdichtung für Turbinenlaufräder
FR2228967A1 (de) * 1973-05-12 1974-12-06 Rolls Royce
FR2293594A1 (fr) * 1974-12-07 1976-07-02 Rolls Royce Perfectionnements aux turbomoteurs
GB1484288A (en) * 1975-12-03 1977-09-01 Rolls Royce Gas turbine engines
US4131388A (en) * 1977-05-26 1978-12-26 United Technologies Corporation Outer air seal
GB2087979A (en) * 1980-11-22 1982-06-03 Rolls Royce Gas turbine engine blade tip seal

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FR1003299A (fr) * 1949-12-13 1952-03-17 Rateau Soc Perfectionnement aux turbines à gaz et autres turbo-machines axiales
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FR2540560B1 (fr) * 1983-02-03 1987-06-12 Snecma Dispositif d'etancheite d'aubages mobiles de turbomachine
FR2548733B1 (fr) * 1983-07-07 1987-07-10 Snecma Dispositif d'etancheite d'aubages mobiles de turbomachine
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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR1219504A (fr) * 1958-03-25 1960-05-18 Zd Y V I Bague de joint étanche pour roue à aubes de turbines à gaz
CH408960A (de) * 1961-09-04 1966-03-15 Licentia Gmbh Radialdichtung für Turbinenlaufräder
FR2228967A1 (de) * 1973-05-12 1974-12-06 Rolls Royce
FR2293594A1 (fr) * 1974-12-07 1976-07-02 Rolls Royce Perfectionnements aux turbomoteurs
GB1484288A (en) * 1975-12-03 1977-09-01 Rolls Royce Gas turbine engines
US4131388A (en) * 1977-05-26 1978-12-26 United Technologies Corporation Outer air seal
GB2087979A (en) * 1980-11-22 1982-06-03 Rolls Royce Gas turbine engine blade tip seal

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB2244524A (en) * 1990-05-31 1991-12-04 Gen Electric Clearance control in gas turbine engines
GB2244524B (en) * 1990-05-31 1994-03-30 Gen Electric Clearance control in gas turbine engines
EP0716220A1 (de) * 1994-12-07 1996-06-12 Societe Nationale D'etude Et De Construction De Moteurs D'aviation "Snecma" Ungeteilter Leitkranz für eine Turbomaschine
FR2728016A1 (fr) * 1994-12-07 1996-06-14 Snecma Distributeur monobloc non-sectorise d'un stator de turbine de turbomachine
EP0770761A1 (de) * 1995-10-23 1997-05-02 United Technologies Corporation Mantelring zur Schaufelspitzenabdichtung

Also Published As

Publication number Publication date
US4787817A (en) 1988-11-29
DE3663556D1 (en) 1989-06-29
EP0192556B1 (de) 1989-05-24
JPS61190101A (ja) 1986-08-23
FR2577281A1 (fr) 1986-08-14
JPH0319883B2 (de) 1991-03-18
FR2577281B1 (fr) 1987-03-20

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