WO1992007180A1 - Enceinte de confinement pour turbine radiale - Google Patents

Enceinte de confinement pour turbine radiale Download PDF

Info

Publication number
WO1992007180A1
WO1992007180A1 PCT/US1991/005794 US9105794W WO9207180A1 WO 1992007180 A1 WO1992007180 A1 WO 1992007180A1 US 9105794 W US9105794 W US 9105794W WO 9207180 A1 WO9207180 A1 WO 9207180A1
Authority
WO
WIPO (PCT)
Prior art keywords
turbine
containment ring
rotor
radially
annular combustor
Prior art date
Application number
PCT/US1991/005794
Other languages
English (en)
Inventor
Phillip Vessa
Jesus Aguilar Garcia
Original Assignee
Sundstrand Corporation
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 Sundstrand Corporation filed Critical Sundstrand Corporation
Publication of WO1992007180A1 publication Critical patent/WO1992007180A1/fr

Links

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
    • F01D21/00—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for
    • F01D21/04—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position
    • F01D21/045—Shutting-down of machines or engines, e.g. in emergency; Regulating, controlling, or safety means not otherwise provided for responsive to undesired position of rotor relative to stator or to breaking-off of a part of the rotor, e.g. indicating such position special arrangements in stators or in rotors dealing with breaking-off of part of rotor

Definitions

  • This invention is generally related to gas turbines and, more particularly, a radial inflow turbine and a containment ring system therefor.
  • containment rings which are generally formed of metal and are disposed radially outwardly of the turbine wheel, usually to some degree axially to each side thereof.
  • the containment rings are made of material with sufficient strength to halt and arrest further radial outward movement upon catastrophic failure of the turbine wheel as to all parts moving radially outwardly due to impingement upon the containment ring.
  • such fragments may damage or destroy such important components as hydraulic, communication, and/or power lines.
  • important components such as hydraulic, communication, and/or power lines.
  • the potential result from this need not be explained.
  • the containment ring or rings usually are brought very near to those areas of the gas turbine that are exposed to high temperatures. This, in turn, means that increasingly exotic materials, whose cost is a consideration, must be used in order to withstand the high temperatures and thermal cycling in the environment in which they are placed. Despite this fact, the containment ring or rings must also be able to reliably provide containment in order to avoid the dangers that are inherent in a catastrophic failure.
  • both fuel and air are injected and ignited in the combustion annulus. It is also conventional for there to be a cooling air film introduced along the walls of the combustion annulus upstream of the dilution zone. Of course, the hot gases that result from combustion then pass toward turbine blades to drive the turbine wheel.
  • An exemplary embodiment of the invention achieves the foregoing objects in a gas turbine which includes a rotor having a radial inflow turbine wheel with turbine blades and a turbine nozzle assembly outwardly of the turbine blades.
  • the nozzle is adapted to direct hot gases at the turbine blades to cause rotation of the rotor.
  • An annular combustor is disposed about the rotor and has an outlet to the nozzle.
  • the annular combustor has spaced inner and outer walls connected by a generally radially extending wall opposite of the outlet.
  • a radial turbine containment ring system includes a first containment ring and a second containment ring. With this arrangement, the first containment ring is outwardly of the rotor and the nozzle and the second containment ring is at least partially about the rotor outwardly of the turbine blades. Additional details of the invention include a housing defining a dilution air flow path about at least a portion of the annular combustor.
  • the dilution air flow path advantageously includes a radially outer compressed air inlet in communication with a compressor supplying dilution air at an upstream end thereof.
  • the first containment ring is fixedly mounted outwardly of the housing at the upstream end of the dilution air flow path.
  • the gas turbine may include a diffuser leading from the compressor to the upstream end of the dilution air flow path.
  • the first, or primary, containment ring is then advantageously fixedly mounted to the diffuser so as to be generally concentric with the outer wall of the annular combustor at a point radially inwardly of the housing, As a result, the first containment ring may advantageously be at least axially coextensive with the turbine blades in a plane of rotation of the rotor.
  • the first containment ring may be fixedly mounted outwardly of the annular combustor at the upstream end of the dilution air flow path. Thus, it may be exposed to cooling air from the compressor that enters the dilution air flow path which means that it may be made of less exotic materials and still have a long life span.
  • the gas turbine preferably includes front and rear turbine shrouds wherein at least the front turbine shroud substantially radially bounds the turbine blades in the plane of rotation thereof.
  • the front turbine shroud preferably includes a radially extending portion and an axially extending portion radially inwardly of the housing.
  • the first containment ring is advantageously disposed between the axially extending portion of the front turbine shroud and the housing.
  • the second containment ring it is advantageously fixedly mounted relative to the housing at a point corresponding to the downstream end of the dilution air flow path.
  • the rear turbine shroud it substantially radially bounds the turbine blades in the plane of rotation thereof and generally conforms to the contour of the turbine blades and extends radially outwardly thereof to the nozzle in a most highly preferred embodiment.
  • the second containment ring is disposed radially outwardly of the rear turbine shroud between the shroud and the combustor at least partially axially coextensive with the turbine blades.
  • a housing substantially surrounds the annular combustor in spaced relation to the inner, outer and radially extending walls thereof including at least the rear turbine wheel shroud.
  • the housing and walls together define a dilution air flow path extending from a compressed air inlet to at least one dilution air outlet where the compressed air inlet is in communication with a source of dilution air.
  • the dilution air outlet is in communication with the annular combustor downstream of the compressed air inlet preferably adjacent the second containment ring.
  • the containment ring system includes a primary containment ring and a secondary containment and deflection ring.
  • the primary containment ring is positioned outwardly of the rotor with the secondary containment and deflection ring being fixedly mounted and generally conforming to the rear turbine wheel shroud for containing fragments within the plane of rotation and deflecting such fragments toward the primary containment ring. While so doing, the containment ring system is arranged so as to be effective for dissipating fragment energy in the event of burst failure.
  • Fig. 1 is a somewhat schematic, fragmentary sectional view of a gas turbine in accordance with the present invention.
  • Fig. 2 is a somewhat schematic, fragmentary sectional view similar to Fig. 1 but illustrating a modified embodiment thereof.
  • the reference numeral 10 designates generally a gas turbine in accordance with the present invention.
  • the gas turbine 10 includes a rotor generally designated 12 which defines a radial inflow turbine wheel 14 having blades 16 and a nozzle 18 outwardly of the turbine blades 16.
  • the nozzle 18 is adapted to direct hot gases at the turbine blades 16 to cause rotation of the rotor 12.
  • the gas turbine 10 also includes an annular combustor 20 about the rotor 12 and having an outlet 22 to the nozzle 18. As shown, the annular combustor 20 has spaced inner and outer walls 24 and 26 connected by a generally radially extending wall 28. Still referring to Fig.
  • a radial turbine containment ring system 32 is defined by a first containment ring 34 and a second containment ring 36.
  • the first containment ring 34 is positioned outwardly of the rotor 12 and the nozzle 18.
  • the second containment ring 36 is at least partially about the rotor 12 immediately outwardly of the turbine blades 16.
  • the gas turbine 10 includes a housing 38 defining a dilution air flow path 40, as shown by the arrows, about at least a portion and preferably all of the annular combustor 20.
  • the dilution air flow path 40 includes a radially outer compressed air inlet as at 42 in communication with a compressor 44 supplying dilution air at an upstream end thereof.
  • the first containment ring 34 is fixedly mounted inwardly of the housing 38 at the upstream end of the dilution air flow path 40 in a plane of rotation of the rotor 12.
  • the gas turbine 10 includes a diffuser 46 leading from the compressor 44 to the upstream end of the dilution air flow path 40.
  • the first containment ring 34 is, in the illustrated embodiment, fixedly mounted to the diffuser 46 (see, also. Fig. 2) by any conventional means and is generally concentric with the outer wall 26 of the annular combustor 20 at a point radially inwardly of the housing 38.
  • the first containment ring 34 is at least axially coextensive with the turbine blades 16 at a point outwardly of the annular combustor 20.
  • Additional details of the gas turbine engine 10 include a pair of front and rear turbine shrouds 48 and 50 wherein the front turbine shroud 48 substantially radially bounds the turbine blades 16.
  • the front turbine shroud 48 radially bounds the turbine blades 16 in the plane of rotation thereof inasmuch as it includes a radially extending portion 48a and an axially extending portion 48b radially inwardly of the housing 38.
  • the first containment ring 34 is suitably disposed between the axially extending portion 48b of the front turbine shroud 48 and the housing 38.
  • the second containment ring 36 is fixedly mounted relative to the housing 38 at a downstream end of the dilution air flow path 40.
  • the rear turbine shroud 50 substantially radially bounds the turbine blades 16 in the plane of rotation thereof and generally conforms to the contour of the turbine blades 16 while extending radially outwardly thereof to the nozzle 18.
  • the second containment ring 36 is disposed radially outwardly of the rear turbine shroud 50 substantially in the plane of rotation.
  • the second containment ring 36 is at least partially axially coextensive with the turbine blades 16 and is preferably disposed between the rear turbine shroud 50 and the annular combustor 20. It is, however, preferably located at a point downstream of the primary combustion zone 52 whereby dilution air flowing about the dilution air flow path 40 may be introduced through openings as at 54 to flow along the combustor-facing surface 36a of the second containment ring 36.
  • the second containment ring 36 also shields a portion of the rear turbine shroud 50 as well as the turbine blades 16 from the heat generated in the annular combustor 20.
  • the housing 38 and walls 24, 26 and 28 together define the dilution air flow path 40.
  • the dilution air flow path 40 extends from the compressed air inlet 42 to at least one dilution air outlet such as 54, and preferably a plurality of such outlets substantially as shown in the drawings whereby air film cooling may be achieved, but it will also be appreciated that a primary dilution air outlet 56 will advantageously be provided for introduction of a large supply of dilution air into a dilution air zone 58 downstream of the combustion zone 52.
  • the dilution air outlets 54 and 56 are all in communication with the annular combustor 20 downstream of the compressed air inlet 42.
  • the containment ring system 32 will be well appreciated as including a primary containment ring 34.
  • the second containment ring 36 essentially comprises a secondary containment and deflection ring inasmuch as its position and shape generally conforming to the rear turbine wheel shroud 50 and turbine blades 16 serves well to not only contain fragments within the plane of rotation but also to deflect such fragments toward the primary containment ring 34.
  • the secondary containment and deflection ' ring 36 also serves to dissipate fragment energy in the event of a burst failure.
  • the secondary containment and deflection ring 36 is fixedly mounted relative to the housing 38 at the downstream end of the dilution air flow path 40 by any conventional means. This may comprise threading the ring 36 to the turbine shroud 50 as at 60, or using pins or any other type of fastening means that will maintain structural integrity even when fragments from the turbine wheel 12 make high energy contact therewith.
  • the secondary containment and deflection ring 36 is suitably disposed at a position which is radially outwardly adjacent the rear turbine shroud 50 but well within in the plane of rotation of the rotor 12.
  • the primary containment ring 34 is mounted so as to be exposed to dilution air at a point which is located externally of the annular combustor 20.
  • the front turbine shroud 48 and the outer wall 26 are each interconnected as at 62 by any conventional means with the primary containment ring 34 being radially outwardly thereof and adjacent the upstream end of the dilution air flow path 40 where it is suitably cooled.
  • the radially extending portion 48a of the front turbine shroud 48 and the radially outermost portion 50a of the rear turbine shroud 50 together define the nozzle 18 radially well inwardly of the primary containment ring 34.
  • the containment ring system 32 serves to dissipate fragment energy as previously described while also serving to reduce fragmentation. The latter occurs due to the close shrouding effects of the primary containment ring 34 and the secondary containment and deflection ring 36 in relation to the turbine blades 16. Still further, the containment ring system 32 serves to deflect and channel fragments within the rotor burst plane minimizing the probability of impact on the housing 38.
  • like reference numerals designate like components in the respective views wherein the gas turbines have many common components. It will be seen, however, that Fig. 2 differs from Fig.
  • the inner wall 24 includes an inner wall segment 24a extending the dilution air flow path 40 whereby the secondary containment and deflection ring 36 is disposed directly in the dilution air flow path outwardly of the annular combustor 20 due to the fact that the inner wall portion 24a extends to a point immediately adjacent the outlet 22 of the combustor 20 where it is secured to the radially outermost portion 50a of the rear turbine shroud 50 adjacent the nozzle 18.
  • this provides even additional cooling effects to the secondary containment and deflection ring 36 but while giving up a certain volume within the dilution zone 58 of the annular combustor 20.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

Afin d'éviter l'enfoncement du carter (38) d'une turbine à gaz (10) qui se produit en cas d'explosion, la turbine à gaz (10) comprend une enceinte de confinement (32). La turbine à gaz (10) est du type comprenant un rotor (12) comportant une roue de turbine (14) pourvue de lames (16) et un ajutage (18) situé à l'extérieur des lames de turbine (16). L'ajutage (18) est conçu pour diriger les gaz brûlants sur les lames de turbine (16) ce qui entraîne la rotation du rotor (12). La turbine à gaz (10) est également du type comprenant un brûleur annulaire (20) situé autour du rotor (12) ainsi qu'un orifice de sortie (22) pour l'ajutage (18). Le brûleur annulaire (20) est défini par des parois espacées internes et externes (24 et 26) reliées par une paroi s'étendant généralement radialement (28). Avec ce système, la turbine à gaz (10) utilise une enceinte de confinement de la turbine radiale (32) qui comprend une première enceinte de confinement (34) située à l'extérieur du rotor (12) et de l'ajutage (18) et une deuxième enceinte de confinement (36) située au moins partiellement autour du rotor (12) à l'extérieur des lames de turbine (16), ce qui permet de renvoyer les fragments vers la première enceinte de confinement (34) tout en dissipant l'énergie cinétique des fragments dans le cas d'une explosion.
PCT/US1991/005794 1990-10-22 1991-08-15 Enceinte de confinement pour turbine radiale WO1992007180A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US60125990A 1990-10-22 1990-10-22
US601,259 1990-10-22

Publications (1)

Publication Number Publication Date
WO1992007180A1 true WO1992007180A1 (fr) 1992-04-30

Family

ID=24406815

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US1991/005794 WO1992007180A1 (fr) 1990-10-22 1991-08-15 Enceinte de confinement pour turbine radiale

Country Status (1)

Country Link
WO (1) WO1992007180A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0718471A1 (fr) * 1994-12-21 1996-06-26 Hispano-Suiza Bouclier de protection d'une turbomachine
WO2002090722A1 (fr) 2001-05-04 2002-11-14 Abb Turbo Systems Ag Dispositif de protection en cas d'eclatement pour compresseur radial de turbocompresseurs
US6533541B1 (en) 2001-12-04 2003-03-18 Honeywell International, Inc. High energy particle arrestor for air turbine starters
US6695574B1 (en) 2002-08-21 2004-02-24 Pratt & Whitney Canada Corp. Energy absorber and deflection device
US9546563B2 (en) 2012-04-05 2017-01-17 General Electric Company Axial turbine with containment shroud

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3652176A (en) * 1970-04-20 1972-03-28 Sudstrand Corp Turbine wheel containment device
US4934899A (en) * 1981-12-21 1990-06-19 United Technologies Corporation Method for containing particles in a rotary machine
US4955192A (en) * 1988-12-12 1990-09-11 Sundstrand Corporation Containment ring for radial inflow turbine

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3652176A (en) * 1970-04-20 1972-03-28 Sudstrand Corp Turbine wheel containment device
US4934899A (en) * 1981-12-21 1990-06-19 United Technologies Corporation Method for containing particles in a rotary machine
US4955192A (en) * 1988-12-12 1990-09-11 Sundstrand Corporation Containment ring for radial inflow turbine

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP0718471A1 (fr) * 1994-12-21 1996-06-26 Hispano-Suiza Bouclier de protection d'une turbomachine
WO1996019641A1 (fr) * 1994-12-21 1996-06-27 Societe Hispano Suiza Bouclier de protection d'une turbomachine
WO2002090722A1 (fr) 2001-05-04 2002-11-14 Abb Turbo Systems Ag Dispositif de protection en cas d'eclatement pour compresseur radial de turbocompresseurs
US6533541B1 (en) 2001-12-04 2003-03-18 Honeywell International, Inc. High energy particle arrestor for air turbine starters
US6814539B2 (en) 2001-12-04 2004-11-09 Honeywell International, Inc. High energy particle arrestor for air turbine starters
US6695574B1 (en) 2002-08-21 2004-02-24 Pratt & Whitney Canada Corp. Energy absorber and deflection device
US9546563B2 (en) 2012-04-05 2017-01-17 General Electric Company Axial turbine with containment shroud

Similar Documents

Publication Publication Date Title
KR100379728B1 (ko) 로터조립체용시라우드및로터조립체시라우드용블레이드외부공기시일
JP6932414B2 (ja) 燃焼器アセンブリ
US5630703A (en) Rotor disk post cooling system
US4700544A (en) Combustors
US6546733B2 (en) Methods and systems for cooling gas turbine engine combustors
US5509270A (en) Gas turbine engine combustor heatshield
US5310319A (en) Free standing turbine disk sideplate assembly
US4948338A (en) Turbine blade with cooled shroud abutment surface
WO1996019640A1 (fr) Assemblage de confinement de moyeu de compresseur centrifuge
US6860110B2 (en) Gas turbine shaft and heat shield cooling arrangement
US4659282A (en) Apparatus for preventing the spreading of titanium fires in gas turbine engines
CN107120685B (zh) 燃烧器组件
JP6584762B2 (ja) ガスタービンのシーリングを容易にする方法およびシステム
CA1311133C (fr) Turbine radiale
EP0205559A1 (fr) Enceinte de protection d'une roue de turbine
US4578942A (en) Gas turbine engine having a minimal blade tip clearance
US11333358B2 (en) One-piece combustion chamber
US4944152A (en) Augmented turbine combustor cooling
US4265590A (en) Cooling air supply arrangement for a gas turbine engine
US20100068069A1 (en) Turbine Blade
US5271220A (en) Combustor heat shield for a turbine containment ring
US4955192A (en) Containment ring for radial inflow turbine
GB1578474A (en) Combustor mounting arrangement
US5129224A (en) Cooling of turbine nozzle containment ring
WO1992007180A1 (fr) Enceinte de confinement pour turbine radiale

Legal Events

Date Code Title Description
AK Designated states

Kind code of ref document: A1

Designated state(s): JP

AL Designated countries for regional patents

Kind code of ref document: A1

Designated state(s): AT BE CH DE DK ES FR GB GR IT LU NL SE