WO2009030197A1 - Anneau de protection multicouche pour système de propulsion d'aéronef - Google Patents

Anneau de protection multicouche pour système de propulsion d'aéronef Download PDF

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Publication number
WO2009030197A1
WO2009030197A1 PCT/DE2008/001417 DE2008001417W WO2009030197A1 WO 2009030197 A1 WO2009030197 A1 WO 2009030197A1 DE 2008001417 W DE2008001417 W DE 2008001417W WO 2009030197 A1 WO2009030197 A1 WO 2009030197A1
Authority
WO
WIPO (PCT)
Prior art keywords
shield
turbine
layers
housing
designed
Prior art date
Application number
PCT/DE2008/001417
Other languages
German (de)
English (en)
Inventor
Wilfried Weidmann
Original Assignee
Mtu Aero Engines Gmbh
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 Mtu Aero Engines Gmbh filed Critical Mtu Aero Engines Gmbh
Priority to CA2698283A priority Critical patent/CA2698283A1/fr
Priority to EP08829278A priority patent/EP2191105A1/fr
Priority to US12/733,382 priority patent/US20100202872A1/en
Publication of WO2009030197A1 publication Critical patent/WO2009030197A1/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
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/14—Casings or housings protecting or supporting assemblies within

Definitions

  • the invention relates to a shielding of a turbine housing of an aircraft engine against radial escape of Schaufeltrüinmern according to the preamble of patent claim 1.
  • Such a shield of the prior art is shown in partial detail in Figure 2.
  • the low-pressure turbine 1 is shown with turbine blades 2, which are arranged within a turbine housing 3.
  • the turbine blades are axially downstream of a compressor, not shown, and a combustion chamber, not shown, and are located on a turbine disk which rotates about the engine axis.
  • the turbine housing 3 is connected via a flange 5 with the turbine exhaust duct 10.
  • the turbine exhaust duct 10 of the prior art is designed as a casting, which also has a containment function due to the existing material thickness. That is, in the unlikely event of engine damage with loss of turbine blades or blade parts, the turbine exhaust port with containment function serves to prevent the escape of the blade parts from the engine housing and thereby possible damage to the Avoid airframe.
  • the decisive for the design impact area is characterized by the angle ⁇ enclosing a straight line.
  • Future engine concepts require low-pressure turbines with high AN 2 , high turbine inlet temperatures and a compact, short design to meet the required specifications.
  • the containment protection is a special design criterion, because due to the higher momentum of loose blade parts, the regular casting thickness of the turbine exhaust duct is no longer sufficient to prevent a possible passage of the blade parts. Therefore, according to the current state of the art, no material, cost and weight optimized low-pressure turbine / turbine exhaust gas duct (LPT / TEC) connection is possible. Rather, the choice of material and material thickness of the LPT / TEC connection is determined by the required containment thickness and not by the optimized LPT / TEC connection. Also, the choice of materials is determined by the higher demands on the casting material in the containment area and thus more expensive.
  • Fiberglass fabric hose is dimensioned so that it fits tightly on the support element.
  • no special solution for the low - pressure turbine / turbine exhaust duct connection is presented here.
  • the disadvantage of this solution is on the one hand the unfixed structure of the collar, which is highly sensitive to external influences, such as mechanical influences, moisture, etc.
  • Another disadvantage is that Damage to the continuous fiber of the fiberglass fabric hose can not be readily noticed and can lead to a total failure of the containment protection in case of need.
  • the invention is therefore based on the object to avoid the disadvantages of the known solutions of the prior art and to provide an improved solution for containment protection on the LPT / TEC connection, in particular of high-speed low-pressure turbines.
  • the shielding according to the invention of a turbine housing of an aircraft engine against radial escape of blade debris, in particular for a high-speed low-pressure turbine is characterized in that the shield is designed as a rigid annular component of several layers.
  • the annular shield may be arranged radially inside or outside the turbine housing. When mounted on the turbine housing can be directed through the shield also targeted cooling air, for example from the fan flow to the outer skin of the housing. Further, it is possible that the annular shield consists of several segments, whereby manufacture and assembly are facilitated. Due to the rigid design, the shield is protected against external influences and can be self-supporting.
  • An advantageous embodiment of the shield according to the invention provides that the shield is arranged on the turbine exhaust duct.
  • the disadvantages of the prior art are avoided.
  • the design of the turbine exhaust duct can be cost and weight optimized, ie, cheaper materials and material thicknesses can be used here than would be the case with integrated containment function.
  • the containment function is then perceived solely by the annular multilayer shield.
  • an advantageous embodiment of the shield according to the invention provides that the shield is designed as a forging component. This allows a multilayer construction with selection of suitable material layers. On the one hand, the strength is the defining factor as well as the temperatures in the area of the low-pressure turbine at the housing or at the LPT / TEC connection. In the case of a circumferentially multi-part shielding ring, the possibility of thermal expansion must be considered.
  • a further advantageous embodiment of the shield according to the invention provides that the shield is arranged within the turbine housing. This prevents on the one hand disturbing structures outside of the turbine housing and on the other hand is prevented by a blade damage the housing or the LPT / TEC connection is penetrated, whereby the cost of a power plant failure continues to rise.
  • the shield is designed as a flow guide. This can be the case both when using the shield inside or outside the housing.
  • additional flow guide can be mounted on the shield or the shield itself is aerodynamically shaped or mounted.
  • the shield is designed as a heat shield. This is particularly necessary when installed in the flow channel, ie within the turbine housing. However, this may also be true when mounted on the outer circumference of the turbine housing be useful to prevent injury from burns on hot engine parts during maintenance.
  • an advantageous embodiment of the shield according to the invention provides that the layers are constructed of different materials. Suitable materials include, for example, malleable high-temperature alloys. As a result, the strength properties, thermal expansion and weight of the shield can be influenced to the desired extent. This makes sense in particular in terms of weight and cost optimization.
  • An advantageous embodiment of the shield according to the invention provides that the layers have different thicknesses. Like the choice of material, the choice of layer thickness can optimize the strength and weight of the shield and thus reduce the cost of the component.
  • An advantageous embodiment of the shield according to the invention provides that the layers are matched to each other in a vibration-optimized manner.
  • the layers of the multi-layer shielding ring are connected without resonance in the shielding housing.
  • both the vibration properties of the shield alone and the vibration characteristics of the components coupled to the shield can be taken into account.
  • an advantageous embodiment of the shield according to the invention provides that the shield has an enclosure for different functional layers. It can fall under the enclosure and a shield housing, with which different layers are connected by joining technology.
  • the annular layers can be bordered or encompassed, as it were, from three sides and may also be accommodated floating within the enclosure. Further measures improving the invention will be described in more detail below together with the description of a preferred exemplary embodiment of the invention with reference to FIGS. Show it:
  • Fig. 2 is a schematic fragmentary view of a shield of the prior art.
  • Directional information refers to the axes of the aircraft propulsion system.
  • FIG. 1 shows a detail schematically an advantageous embodiment of a shield 6 according to the invention on a high-speed low-pressure turbine 1.
  • the compressor not shown in the drawing and the combustion chamber, and also not shown high and medium pressure turbine.
  • FIG. 1 shows a section of a half section.
  • FIG. 1 shows a part of a turbine blade 2, which is arranged within a turbine housing 3 surrounding the turbine stage in the circumferential direction.
  • the turbine housing 3 is connected via a material-technically optimized flange 5 with the turbine exhaust duct 4 and connected thereto.
  • the shield 6 is arranged at the connection of the low-pressure turbine 1 to the turbine exhaust duct 4 within the turbine housing 3.
  • a flange 9 protrudes, on which the shield 6 and the shield housing 7 is flanged.
  • the shield L-shaped and annular in the circumferential direction 6 or the containment ring is designed as a multilayer forging.
  • the two layers 8 of the shield 6 are accommodated in a shielding housing 7 and connected to this forging technology. Both the type of alloy and the layer thickness / number of layers differ in the case of the two layers 8 shown in the exemplary embodiment.
  • the resonance-free shielding 6 has, in addition to the containment function, also integrated heat shield and flow conduction function.
  • the containment function is not integrated in the connection between low-pressure turbine 1 and turbine exhaust gas channel 4, as a result of which this connection can be configured as a weight-optimized casting.

Landscapes

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

Abstract

L'invention concerne un bouclier (6) d'un carter de turbine (3) d'un groupe motopropulseur empêchant la sortie radiale de débris d'aubes, notamment pour une turbine basse pression à régime rapide. L'invention se caractérise en ce que le bouclier (6) est réalisé sous la forme d'un composant annulaire rigide, constitué de plusieurs couches (8). L'invention pallie les inconvénients de l'état de la technique grâce à une conception du canal d'échappement de la turbine, réalisé sous forme de pièce coulée, où la fonction de protection est dissociée. En particulier, la conception du canal d'échappement de la turbine peut être optimisée en termes de coût et de poids.
PCT/DE2008/001417 2007-09-07 2008-08-27 Anneau de protection multicouche pour système de propulsion d'aéronef WO2009030197A1 (fr)

Priority Applications (3)

Application Number Priority Date Filing Date Title
CA2698283A CA2698283A1 (fr) 2007-09-07 2008-08-27 Anneau de protection multicouche pour systeme de propulsion d'aeronef
EP08829278A EP2191105A1 (fr) 2007-09-07 2008-08-27 Anneau de protection multicouche pour système de propulsion d'aéronef
US12/733,382 US20100202872A1 (en) 2007-09-07 2008-08-27 Multilayer shielding ring for a flight driving mechanism

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE102007042767.2 2007-09-07
DE102007042767A DE102007042767A1 (de) 2007-09-07 2007-09-07 Mehrschichtiger Abschirmungsring für einen Flugantrieb

Publications (1)

Publication Number Publication Date
WO2009030197A1 true WO2009030197A1 (fr) 2009-03-12

Family

ID=40221258

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/DE2008/001417 WO2009030197A1 (fr) 2007-09-07 2008-08-27 Anneau de protection multicouche pour système de propulsion d'aéronef

Country Status (5)

Country Link
US (1) US20100202872A1 (fr)
EP (1) EP2191105A1 (fr)
CA (1) CA2698283A1 (fr)
DE (1) DE102007042767A1 (fr)
WO (1) WO2009030197A1 (fr)

Families Citing this family (29)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10167779B2 (en) * 2012-09-28 2019-01-01 United Technologies Corporation Mid-turbine frame heat shield
US9631517B2 (en) 2012-12-29 2017-04-25 United Technologies Corporation Multi-piece fairing for monolithic turbine exhaust case
US9982564B2 (en) 2012-12-29 2018-05-29 United Technologies Corporation Turbine frame assembly and method of designing turbine frame assembly
WO2014105826A1 (fr) 2012-12-29 2014-07-03 United Technologies Corporation Disque et ensemble de support d'étanchéité
WO2014105619A1 (fr) 2012-12-29 2014-07-03 United Technologies Corporation Bossage multifonction pour carter de sortie turbine
US10294819B2 (en) 2012-12-29 2019-05-21 United Technologies Corporation Multi-piece heat shield
US10006306B2 (en) 2012-12-29 2018-06-26 United Technologies Corporation Turbine exhaust case architecture
WO2014143329A2 (fr) 2012-12-29 2014-09-18 United Technologies Corporation Trous de refroidissement pour jonction de châssis
US9828867B2 (en) 2012-12-29 2017-11-28 United Technologies Corporation Bumper for seals in a turbine exhaust case
US9850774B2 (en) 2012-12-29 2017-12-26 United Technologies Corporation Flow diverter element and assembly
US10087843B2 (en) 2012-12-29 2018-10-02 United Technologies Corporation Mount with deflectable tabs
WO2014105512A1 (fr) 2012-12-29 2014-07-03 United Technologies Corporation Liaison mécanique destinée à un écran thermique segmenté
US9903216B2 (en) 2012-12-29 2018-02-27 United Technologies Corporation Gas turbine seal assembly and seal support
US10472987B2 (en) 2012-12-29 2019-11-12 United Technologies Corporation Heat shield for a casing
US10240481B2 (en) 2012-12-29 2019-03-26 United Technologies Corporation Angled cut to direct radiative heat load
US9903224B2 (en) 2012-12-29 2018-02-27 United Technologies Corporation Scupper channelling in gas turbine modules
US9845695B2 (en) 2012-12-29 2017-12-19 United Technologies Corporation Gas turbine seal assembly and seal support
WO2014105780A1 (fr) 2012-12-29 2014-07-03 United Technologies Corporation Ensemble et support de joint de turbine à gaz à usages multiples
EP2938857B2 (fr) 2012-12-29 2020-11-25 United Technologies Corporation Bouclier thermique pour le refroidissement d'une entretoise
US10138742B2 (en) 2012-12-29 2018-11-27 United Technologies Corporation Multi-ply finger seal
DE112013006325T5 (de) 2012-12-31 2015-11-19 United Technologies Corporation Mehrteiliger Rahmen eines Turbinenabgasgehäuses
EP2938860B1 (fr) 2012-12-31 2018-08-29 United Technologies Corporation Cadre à multiples pièces de compartiment d'échappement de turbine
GB2524220B (en) 2012-12-31 2020-05-20 United Technologies Corp Turbine exhaust case multi-piece frame
US10330011B2 (en) 2013-03-11 2019-06-25 United Technologies Corporation Bench aft sub-assembly for turbine exhaust case fairing
DE102013214389A1 (de) 2013-07-23 2015-01-29 MTU Aero Engines AG Gehäusecontainment
DE102014208883A1 (de) * 2014-05-12 2015-12-03 MTU Aero Engines AG Verfahren zum Auslegen einer Turbine
FR3054527B1 (fr) * 2016-07-29 2019-08-30 Airbus Operations Ensemble pour aeronef comprenant un bouclier de protection contre un eclatement moteur, monte sur le carter d'un module de turbomachine
US10550718B2 (en) 2017-03-31 2020-02-04 The Boeing Company Gas turbine engine fan blade containment systems
US10487684B2 (en) 2017-03-31 2019-11-26 The Boeing Company Gas turbine engine fan blade containment systems

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US3097824A (en) * 1958-11-26 1963-07-16 Bendix Corp Turbine, wheel containment
US3241813A (en) * 1964-01-21 1966-03-22 Garrett Corp Turbine wheel burst containment means
GB1245415A (en) * 1968-09-13 1971-09-08 Rolls Royce Improvements in or relating to fluid flow machines
US3849022A (en) * 1973-07-12 1974-11-19 Gen Motors Corp Turbine blade coolant distributor
DE7501892U (de) * 1975-01-23 1976-06-03 Motoren- Und Turbinen-Union Muenchen Gmbh, 8000 Muenchen Metall-keramik-heissgasfuehrung mit berstschutz-eigenschaften
EP0027756A1 (fr) * 1979-10-19 1981-04-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Dispositif de sécurité en cas de rupture d'élément rotatif de turbomachine
US4547122A (en) * 1983-10-14 1985-10-15 Aeronautical Research Associates Of Princeton, Inc. Method of containing fractured turbine blade fragments
DE4223496A1 (de) * 1992-07-17 1994-01-20 Asea Brown Boveri Vorrichtung zum Reduzieren der kinetischen Energie von berstenden Teilen
WO1999054598A1 (fr) * 1998-04-20 1999-10-28 Pratt & Whitney Canada Corp. Systeme de confinement permettant de maitriser la rupture d'aubes
US20060233636A1 (en) * 2002-06-05 2006-10-19 Volvo Aero Corporation Turbine and a component

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Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3097824A (en) * 1958-11-26 1963-07-16 Bendix Corp Turbine, wheel containment
US3241813A (en) * 1964-01-21 1966-03-22 Garrett Corp Turbine wheel burst containment means
GB1245415A (en) * 1968-09-13 1971-09-08 Rolls Royce Improvements in or relating to fluid flow machines
US3849022A (en) * 1973-07-12 1974-11-19 Gen Motors Corp Turbine blade coolant distributor
DE7501892U (de) * 1975-01-23 1976-06-03 Motoren- Und Turbinen-Union Muenchen Gmbh, 8000 Muenchen Metall-keramik-heissgasfuehrung mit berstschutz-eigenschaften
EP0027756A1 (fr) * 1979-10-19 1981-04-29 Societe Nationale D'etude Et De Construction De Moteurs D'aviation, "S.N.E.C.M.A." Dispositif de sécurité en cas de rupture d'élément rotatif de turbomachine
US4547122A (en) * 1983-10-14 1985-10-15 Aeronautical Research Associates Of Princeton, Inc. Method of containing fractured turbine blade fragments
DE4223496A1 (de) * 1992-07-17 1994-01-20 Asea Brown Boveri Vorrichtung zum Reduzieren der kinetischen Energie von berstenden Teilen
WO1999054598A1 (fr) * 1998-04-20 1999-10-28 Pratt & Whitney Canada Corp. Systeme de confinement permettant de maitriser la rupture d'aubes
US20060233636A1 (en) * 2002-06-05 2006-10-19 Volvo Aero Corporation Turbine and a component

Also Published As

Publication number Publication date
EP2191105A1 (fr) 2010-06-02
US20100202872A1 (en) 2010-08-12
CA2698283A1 (fr) 2009-03-12
DE102007042767A1 (de) 2009-03-12

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