EP0081405A1 - Ringförmige luftgekühlte abreissbare Schaufeldichtung für eine Gasturbine oder einen Kompressor - Google Patents

Ringförmige luftgekühlte abreissbare Schaufeldichtung für eine Gasturbine oder einen Kompressor Download PDF

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Publication number
EP0081405A1
EP0081405A1 EP82402064A EP82402064A EP0081405A1 EP 0081405 A1 EP0081405 A1 EP 0081405A1 EP 82402064 A EP82402064 A EP 82402064A EP 82402064 A EP82402064 A EP 82402064A EP 0081405 A1 EP0081405 A1 EP 0081405A1
Authority
EP
European Patent Office
Prior art keywords
channels
annular
zone
wear
ring
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
EP82402064A
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English (en)
French (fr)
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EP0081405B1 (de
Inventor
Christian Bernard Aubert
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
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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 EP0081405A1 publication Critical patent/EP0081405A1/de
Application granted granted Critical
Publication of EP0081405B1 publication Critical patent/EP0081405B1/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
    • F01D11/12—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator using a rubstrip, e.g. erodible. deformable or resiliently-biased part
    • 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
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling

Definitions

  • the material of the wear layer is most often a porous material (aggregate, felt, foam, perforated plate, etc.) in order to allow its abrasion by the ends of the blades. If no particular provision is made, it is therefore traversed by the cooling air, the flow of which flows at least in part towards the stream of hot gases.
  • the device of the invention makes it possible to avoid the same drawbacks and provides the same advantages, but its structure is simpler and its realization is easier.
  • the invention therefore makes it possible to obtain, by very simple means, a rigorous separation of the functions of the two zones.
  • the radial thermal gradient in the cooling zone is very low since its channels are cooled.
  • it is very high in the wear zone and causes differential expansions there which favor the propagation of the incipient fractures that are micro-cracks.
  • the seal ring being advantageously brazed in the support ring and having to be made of refractory material, it is advantageous to choose to constitute it a superalloy, that is to say an alloy comprising in weight content, more than 50% nickel and / or cobalt.
  • the most suitable method for piercing channels in this material is by electron bombardment. It in fact causes, as the hole progresses to form a channel, intense local heating followed by rapid cooling by diffusion of heat into the mass of the ring.
  • the second solution is more advantageous because it allows the cooling air flow rate to be adjusted as best as possible by passing through the various elementary seal ring devices "in parallel".
  • the turbine ring 10 which surrounds the wheel, the end E of which can be seen from a blade shown in broken lines, can for example be interposed between an outer distributor ring of the stage in question and , if necessary, an external distributor ring for the next stage.
  • the cooled seal device comprises the support ring 20 and the seal ring 30.
  • the support ring 20 is fixed at its ends to the turbine ring 10 by means of circular weld beads 21 and it is also centered if necessary by ribs 11.
  • the seal ring 30 is housed in the support ring 20 to which it is brazed by its periphery 31. It is traversed by a plurality of parallel channels 32 which are represented in FIG. 1 by dashed lines, some of which are seen in section in Figure 2 and occupying its entire section.
  • An annular surface 22 belonging to an upstream flange 23 of the support ring 20 is brazed on the upstream face 34 of the seal ring 30 while a ring 33 can be brazed on the downstream face 35 of the same ring.
  • the internal circumferences of this bearing surface and of this ring are flush with the internal contour 36 of this ring 30, of radius R1, while that their external circumferences have equal radii R3 substantially smaller than the radius R2 of the external contour 31.
  • Said bearing surface and said ring therefore form screens which divide the ring 30 into two concentric annular zones, namely on the one hand a zone external Zl of external radius R2 and internal radius R3 and on the other hand an internal zone Z2 of external diameter R3 and internal diameter Rl. These screens transform the channels 32 located in the zone Z2 into closed or half-open cavities, if the ring 33 is not brazed.
  • An air flow obtained by bypassing a fraction of the flow rate of the compressor supplying the turbine first enters through a plurality of orifices 12 (formed in the ring 10) in the annular chamber 13 which surrounds the upstream part of the support ring 20, then it penetrates by means of orifices 24 formed therein in an annular chamber 25 delimited by the upstream face of the seal ring 30 and by the flange 23; it is then blown into the channels 32 of the zone Z1 and exits through the downstream face of said zone.
  • the ring 30 is constituted by the stack of elementary rings 37, drilled identically and mounted so that the channels 32 are perfectly aligned.
  • the wear zone Z2 the channels of which constitute closed cavities, is thermally insulating and the radial temperature gradient in operation is large.
  • Figure 3 shows an embodiment which eliminates these constraints.
  • the refractory wear ring is divided into elementary rings of short length 67 each of which is provided with its own means for supplying cooling air.
  • the turbine ring 40 has as many rows of air passage openings 42 as there are elementary rings 67 and the support ring is divided into as many support elements 56 each of which houses a ring elementary 67 which is brazed there by its periphery.
  • Each element 56 is provided with an upstream internal flange 57 on which this elementary ring abuts and which is shaped so as to provide an annular chamber 55 opposite the zone Zl (see FIG. 2).
  • each elementary ring 67 is shorter than the housing reserved for it in the corresponding element 56, which provides a vacuum 58 between the downstream end of this elementary ring and the flange 57 which follows.
  • the channels of zone Z2 of each elementary ring are closed using screens annular 62 brazed on the upstream end of each ring.
  • An annular screen 63 can also be brazed on the downstream ends of each ring.
  • the rows of openings 42 are separated by ribs 41 each of which supports the downstream end of an element 56 and the upstream end of the element which follows it and which delimit annular chambers 43. Each of these is supplied by the corresponding row of orifices 42 and communicates with the corresponding annular chamber 55 by a row of openings 54 formed in the corresponding element 56.
  • Two annular weld beads 51 secure the stack of support elements 56 flush with the upstream end and the downstream end of the turbine ring 40.
  • the joint device of FIG. 3 in fact consists of the stacking of elementary devices, each of which is practically in accordance with FIG. 1 but which are short enough so that it is not necessary to fragment their joint rings 67. It allows furthermore to admit, at equal supply pressure, a much greater cooling air flow than the device in FIG. 1 since the number of intake openings and circulation channels is much higher while the channels are much shorter. Conversely, to obtain the same air flow, the air pressure required is much lower. It may further be noted that, with the exception of the seal element 67 on the left, each of those which follow it has its internal contour cooled by the film of air delivered by the annular chamber 58 which precedes.
  • FIG 4 illustrates an alternative embodiment of the air intake chamber in the channels of the zone Zl (25, Figure 1; 55, Figure 3).
  • the annular flange 71 (which acts as an abutment for the flanges 23 or 57 of Figures 1 or 3) is planar.
  • the air intake chamber 72 is obtained by contouring the seal ring 73 to obtain an annular recess limited by the rays R2 and R3 (zone Zl) and is supplied with cooling air by openings 74 drilled in the joint support ring 75.
  • the ring 73 is brazed by its non-contoured part (zone Z2) on the flange 71 which therefore plays not only the role of stop but also that of shutter.
  • This variant can be of great interest, in particular if the stage considered is a high pressure compressor stage, because it makes it possible to take an air flow for cooling this stage to a low pressure stage when this cooling would be impossible if this flow should return to the high pressure stream since there would be reversal of the direction of flow.
  • the channels 32 of the zone Z1 and those of the zone Z2 being different, they can be given different diameters and even different relative arrangements.
  • the diameter and pitch of these channels depend on the pressure of Ali "air mentation, pressure to overcome in the vein (if the air must return it) and of the flow required to obtain effective cooling.
  • the diameter must not fall below a certain value to limit the pressure losses and the risk of blockage by dust.
  • 1 mm channels can be provided with a pitch of 1.5 mm.
  • zone Z2 wear zone
  • the channels must be as close as possible and of sufficiently small diameter, and preferably distributed in staggered rows to improve their machinability by the blade tips in the event of of friction, and ensuring a sufficient and homogeneous radial temperature gradient. It is then possible, for example, to provide in this zone channels with a diameter of 0.3 mm arranged in circular rows, the pitch of the channels in each row being 0, 4 mm and these rows being shifted from one to the other by a value equal to half a step so that a determined channel is equidistant from all its neighbors.
  • the closure of the channels of the zone Z2 can be ensured by means of a simple application of solder instead of being soldered by means of a flange or a screen.
  • the seal 30 (or the stack of seals 67) had a conical shape (instead of cylindrical) in the case where the blade ends generate in their movement a conical surface instead of a cylindrical surface as shown in the accompanying drawings.
  • the direction of the channels 32 must, in this case, be parallel to the generatrices of the cone instead of being parallel to the axis of the wheel.

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)
EP82402064A 1981-11-16 1982-11-10 Ringförmige luftgekühlte abreissbare Schaufeldichtung für eine Gasturbine oder einen Kompressor Expired EP0081405B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8121353A FR2516597A1 (fr) 1981-11-16 1981-11-16 Dispositif annulaire de joint d'usure et d'etancheite refroidi par l'air pour aubage de roue de turbine a gaz ou de compresseur
FR8121353 1981-11-16

Publications (2)

Publication Number Publication Date
EP0081405A1 true EP0081405A1 (de) 1983-06-15
EP0081405B1 EP0081405B1 (de) 1985-04-24

Family

ID=9264013

Family Applications (1)

Application Number Title Priority Date Filing Date
EP82402064A Expired EP0081405B1 (de) 1981-11-16 1982-11-10 Ringförmige luftgekühlte abreissbare Schaufeldichtung für eine Gasturbine oder einen Kompressor

Country Status (6)

Country Link
US (1) US4468168A (de)
EP (1) EP0081405B1 (de)
JP (1) JPS58135306A (de)
CA (1) CA1198374A (de)
DE (1) DE3263299D1 (de)
FR (1) FR2516597A1 (de)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0495256A1 (de) * 1991-01-14 1992-07-22 General Motors Corporation Gasturbinendeckband

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4626169A (en) * 1983-12-13 1986-12-02 United Technologies Corporation Seal means for a blade attachment slot of a rotor assembly
FR2576637B1 (fr) * 1985-01-30 1988-11-18 Snecma Anneau de turbine a gaz.
FR2857406B1 (fr) * 2003-07-10 2005-09-30 Snecma Moteurs Refroidissement des anneaux de turbine
US7018113B1 (en) 2003-11-18 2006-03-28 Optiworks, Inc. Optical module package
FR2876933B1 (fr) * 2004-10-25 2008-05-09 Snecma Moteurs Sa Buse pour tete de percage ou d'usinage par faisceau laser
US8408304B2 (en) * 2008-03-28 2013-04-02 Baker Hughes Incorporated Pump mechanism for cooling of rotary bearings in drilling tools and method of use thereof
US8444371B2 (en) * 2010-04-09 2013-05-21 General Electric Company Axially-oriented cellular seal structure for turbine shrouds and related method
US9074597B2 (en) 2011-04-11 2015-07-07 Baker Hughes Incorporated Runner with integral impellor pump
US9181877B2 (en) 2012-09-27 2015-11-10 United Technologies Corporation Seal hook mount structure with overlapped coating
US10197069B2 (en) * 2015-11-20 2019-02-05 United Technologies Corporation Outer airseal for gas turbine engine
US10443426B2 (en) * 2015-12-17 2019-10-15 United Technologies Corporation Blade outer air seal with integrated air shield

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3425665A (en) * 1966-02-24 1969-02-04 Curtiss Wright Corp Gas turbine rotor blade shroud
US3719365A (en) * 1971-10-18 1973-03-06 Gen Motors Corp Seal structure
GB1308771A (en) * 1966-11-02 1973-03-07 Gen Electric Fluid cooled porous stator structure
US3825364A (en) * 1972-06-09 1974-07-23 Gen Electric Porous abradable turbine shroud
GB2062115A (en) * 1979-10-12 1981-05-20 Gen Electric Method of constructing a turbine shroud

Family Cites Families (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3213789A (en) * 1963-10-30 1965-10-26 Braco Engineering Company Method of making rubber printing plates
US3970319A (en) * 1972-11-17 1976-07-20 General Motors Corporation Seal structure
US3893786A (en) * 1973-06-07 1975-07-08 Ford Motor Co Air cooled shroud for a gas turbine engine
US4130373A (en) * 1976-11-15 1978-12-19 General Electric Company Erosion suppression for liquid-cooled gas turbines
FR2393994A1 (fr) * 1977-06-08 1979-01-05 Snecma Materiau abradable metallique et son procede de realisation
FR2401310A1 (fr) * 1977-08-26 1979-03-23 Snecma Carter de turbine de moteur a reaction
FR2468741A1 (fr) * 1979-10-26 1981-05-08 Snecma Perfectionnements aux anneaux a joint d'etancheite refroidi par l'air pour roues de turbine a gaz

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3425665A (en) * 1966-02-24 1969-02-04 Curtiss Wright Corp Gas turbine rotor blade shroud
GB1308771A (en) * 1966-11-02 1973-03-07 Gen Electric Fluid cooled porous stator structure
US3719365A (en) * 1971-10-18 1973-03-06 Gen Motors Corp Seal structure
US3825364A (en) * 1972-06-09 1974-07-23 Gen Electric Porous abradable turbine shroud
GB2062115A (en) * 1979-10-12 1981-05-20 Gen Electric Method of constructing a turbine shroud

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0495256A1 (de) * 1991-01-14 1992-07-22 General Motors Corporation Gasturbinendeckband

Also Published As

Publication number Publication date
EP0081405B1 (de) 1985-04-24
FR2516597B1 (de) 1984-05-11
DE3263299D1 (en) 1985-05-30
CA1198374A (fr) 1985-12-24
US4468168A (en) 1984-08-28
JPS58135306A (ja) 1983-08-11
JPS6313004B2 (de) 1988-03-23
FR2516597A1 (fr) 1983-05-20

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