EP2354464B1 - Gegossene Ummantelungsschlitze mit Vorwirbelleck - Google Patents

Gegossene Ummantelungsschlitze mit Vorwirbelleck Download PDF

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Publication number
EP2354464B1
EP2354464B1 EP11151690.2A EP11151690A EP2354464B1 EP 2354464 B1 EP2354464 B1 EP 2354464B1 EP 11151690 A EP11151690 A EP 11151690A EP 2354464 B1 EP2354464 B1 EP 2354464B1
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EP
European Patent Office
Prior art keywords
slot
turbine engine
core assembly
shroud
refractory
Prior art date
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Application number
EP11151690.2A
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English (en)
French (fr)
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EP2354464A2 (de
EP2354464A3 (de
Inventor
Michael G. Mccaffrey
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.)
RTX Corp
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United Technologies Corp
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Publication date
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Publication of EP2354464A3 publication Critical patent/EP2354464A3/de
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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/02—Sand moulds or like moulds for shaped castings
    • B22C9/04—Use of lost patterns
    • 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
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F01D5/225—Blade-to-blade connections, e.g. for damping vibrations by shrouding
    • 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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • 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
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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
    • F05D2240/55—Seals
    • F05D2240/57—Leaf seals

Definitions

  • the present disclosure is directed to a refractory core assembly for forming cast shroud slots with pre-swirled leakage, a process for casting turbine engine components, such as turbine vane shrouds, having as-cast shroud slots with pre-swirled leakage, and to a cast turbine engine component having shroud slots for providing pre-swirled leakage.
  • JP S60 22002A discloses a blade structure for preventing secondary flow.
  • EP 1096108 A2 discloses a shroud assembly in which the edges of adjacent shroud segments are shaped so as to define a non-linear curvature that matches the curvature of gas flow through the turbine rotor, in order to reduce boundary layer tripping.
  • a system of elements configured to interlock as a refractory metal core assembly configurable in a die for forming a seal slot having a curved trailing edge portion in a turbine vane shroud assembly, said system comprising a first element having a longitudinally extending slot and a second element having a longitudinally extending slot for receiving a portion of said first element so that the first element and the second element interlock, characterised in that said second element has a planar portion and a non-planar portion comprising a curved portion adjacent said planar portion, wherein said curved portion is configured in a shape to form a curved trailing edge portion of a said seal slot.
  • a process for forming a turbine engine component comprising a turbine vane shroud, the process comprising the steps of: placing a refractory core assembly comprising a first element and a second element joined to said first element in a die, the first element having a longitudinally extending slot and the second element having a longitudinally extending slot receiving a portion of the first element so that the first element and the second element interlock, the refractory core assembly being configured to form a seal slot in a said turbine vane shroud; encapsulating said refractory core assembly in a wax pattern having the form of said turbine engine component; forming a ceramic shell mold about said wax pattern; removing said wax pattern; and pouring molten material into said ceramic shell mold to form said turbine engine component; characterised in that: said second element has a planar portion and a non-planar portion comprising a curved portion adjacent said planar portion.
  • Fig. 1 illustrates a typical turbine engine component 10 having a plurality of airfoils 12 and a segmented shroud ring 14 having a linear slot 16 between two of the airfoils 12.
  • the angle of the slot 16 is constrained by the geometry of the airfoils 12. Steep slot angles are not possible due to the airfoil geometry.
  • Fig. 2 illustrates a turbine engine component 10' fabricated in accordance with the present invention.
  • the turbine engine component 10' has a segmented shroud ring 14' with a plurality of airfoils 12'.
  • the slot 16' is not linear.
  • the slot 16' has a contour which is curved to turn leak flow 18' to match rotor motion and the flow of fluid around the turbine engine component 10' as exemplified by the arrow 20'.
  • One or more as-cast slots 16' may be produced in a wall of a shroud ring 14' in between two of the airfoils 12'. Each slot 16' may be cast integrally using the metal refractory core system 30' shown in Figs. 3 - 5 .
  • the refractory metal core system 30' is formed from two thin plates 32' and 34'.
  • the thin plates 32' and 34' are constructed so that they can be interlocked perpendicular to each other.
  • the plate 32' may have a planar construction with two opposed surfaces 130' and 132'.
  • the plate 32' further may have a longitudinally extending slot 36' for receiving the plate 34'.
  • the plate 32' may have a length which is shorter than the length of the plate longitudinally extending 34'. While the plate 32' has been described as having a planar, it may also be non-planar if desired.
  • the plate 34' has a planar portion 134' with opposed surfaces 136' and 138'. Still further, the plate 34' has a longitudinally extending slot 38' extending from a leading edge 40', which slot receives a planar portion of the plate 32'.
  • the plate 34' has a curved trailing edge portion 42' in the shape of the desired configuration for the trailing edge portion 42' of the slot 16'. The exact curvature of the trailing edge portion 42' and the angle are determined by the swirl needs and any airfoil limitations. If desired, the trailing edge portion 42' may also have a twist portion 144' to tailor the radial swirl.
  • Each of the plates 32' and 34' may be formed from any refractory metal or refractory metal alloy. While the plates 32' and 34' may be formed from molybdenum or a molybdenum alloy, they could also be formed from any other suitable refractory material. If desired, each plate 32' and 34' may have a thin ceramic coating applied to the base refractory metal, refractory metal alloy, or refractory material forming the respective plate.
  • the plates 32' and 34' When assembled, as shown in Figs. 4 and 5 , the plates 32' and 34' form an assembly in which the surfaces 130' and 132' are at an angle with respect to surfaces 136' and 138' and the plates 32' and 34' are interlocked.
  • the surfaces 130' and 132' may be perpendicular to the surfaces 136' and 138'.
  • the turbine engine component 10' including the airfoils 12', and the shroud ring 14' may be formed using any suitable technique known in the art.
  • the refractory core assembly comprising a first planar element, namely plate 32', and a second non-planar element, namely plate 34', joined to the first planar element may be placed in a die.
  • the refractory core assembly is encapsulated in a wax pattern having the form of the turbine engine component 10'.
  • a ceramic shell mold is formed about the wax pattern.
  • the wax pattern is removed.
  • step 210 molten material is poured into the ceramic shell mold to form said turbine engine component.
  • step 212 the airfoils and the turbine engine component may be removed from the ceramic shell mold.
  • step 214 the refractory core assembly is removed after said molten material has solidified so as to form an as-cast slot in a shroud portion of the turbine engine component.
  • the refractory core assembly removal step may be performed using an acid leach operation.
  • the refractory core assembly placing step preferably comprises placing a refractory core assembly wherein said first element is fitted into a slot in said second non-planar element and said second non-planar element is fitted into a slot in said planar element.
  • a plurality of refractory core assemblies may be placed in the die. Further, each refractory core assembly may be placed in a portion of the die to be used to form an outer shroud ring and/or an inner shroud ring.
  • the molten material pouring step may comprise pouring a molten material into said die to form a plurality of airfoils, such as pouring a nickel based superalloy.
  • the resultant turbine engine component 10' formed by the foregoing process has an as-cast shroud 14' having a tailored segmentation in the form of an integral feather seal slot 16' with a curved trailing edge portion 50'.
  • Fig. 7 is a sectional view taken along lines 7 - 7 in Fig. 6 and shows the internal portion 52' of the slot 16' formed by the plate 32'.
  • the curved trailing edge portion 52' alters the leakage flow so that it is more aligned with the rotating flow.
  • the slot 16' may have a linear portion 53' adjacent a first edge 56' and the curved or angled portion 50' adjacent a second edge 58'.
  • Fig. 8 illustrates the segmented shroud ring 14' with a flow of fluid 54' being discharged from the slot 16'.
  • the fluid is pre-swirled and flows in the direction of rotation. Such a flow reduces losses.
  • Separation (segmentation) of the shroud is useful because it relieves stress caused by the ring-strut-ring structure, i.e. a hot airfoil is overly constrained by cold inner and outer diameter rings.
  • refractory metal core plates such as plates 32' and 34' is advantageous because one can form complex slot shapes directly into the casting, without complex machining operations. The more complex the shape, the less likely the slot can be machined.
  • the refractory metal core plates can be made very thin compared to a ceramic core.
  • a ceramic core of similar thickness i.e. 0.008 to 0.010" (0.203 to 0.254 mm) or less, would likely result in low casting yield because it could easily break during handling, assembly, wax injection, or during the pour of molten metal. Ceramics are very fragile compared to metals.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Molds, Cores, And Manufacturing Methods Thereof (AREA)

Claims (11)

  1. System (30') aus Elementen, die dazu konfiguriert sind, als eine hochschmelzende Metallkernanordnung ineinanderzugreifen, die in einer Form konfigurierbar ist, um einen Dichtungsschlitz (16') herzustellen, der einen gekrümmten Hinterkantenabschnitt in einer Turbinenschaufelummantelungsanordnung aufweist, wobei das System ein erstes Element (32'), das einen in Längsrichtung verlaufenden Schlitz (36') aufweist, und ein zweites Element (34') umfasst, das einen in Längsrichtung verlaufenden Schlitz (38') zum Aufnehmen eines Abschnitts des ersten Elements (32') aufweist, so dass das erste Element (32') und das zweite Element (34') ineinandergreifen, dadurch gekennzeichnet, dass das zweite Element (34') einen planaren Abschnitt (134') und einen nichtplanaren Abschnitt (42') aufweist, der einen gekrümmten Abschnitt angrenzend an den planaren Abschnitt (134') umfasst, wobei der gekrümmte Abschnitt (42') in einer solchen Form konfiguriert ist, dass er einen gekrümmten Hinterkantenabschnitt des Dichtungsschlitzes (16') bildet.
  2. System nach Anspruch 1, wobei das erste Element (32') ein Paar gegenüberliegender Flächen aufweist und das zweite Element (34') ein Paar gegenüberliegender Flächen aufweist, die in einem Winkel in Bezug auf die Flächen des ersten Elements (32') ausgerichtet sind.
  3. System nach Anspruch 1 oder 2, wobei der gekrümmte Abschnitt (42') dazu konfiguriert ist, einen Hinterkantenabschnitt eines Dichtungsschlitzes (16') in einer solchen Form zu bilden, dass eine Leckströmung, die aus dem Dichtungsschlitz (16') in einen segmentierten Schaufelummantelungsring des Gasturbinentriebwerks ausgetragen wird, im Vergleich zu einer Strömung, die in diesen Dichtungsschlitz (16') eintritt, mehr auf eine rotierende Strömung des Gasturbinentriebwerks neben der Turbinenschaufelummantelung ausgerichtet ist.
  4. System nach einem der vorstehenden Ansprüche, wobei das erste Element (32') eine erste Länge aufweist und das zweite Element (34') eine zweite Länge aufweist, die größer als die erste Länge ist.
  5. System nach einem der vorstehenden Ansprüche, wobei jedes aus dem ersten und dem zweiten Element (32', 34') aus einem hochschmelzenden Metallmaterial hergestellt ist.
  6. System nach einem der vorstehenden Ansprüche, wobei das zweite Element (34') einen Abschnitt (144') mit einer Verdrehung aufweist.
  7. Verfahren zum Herstellen einer Turbinentriebwerkskomponente, die eine Turbinenschaufelummantelung umfasst, wobei das Verfahren die folgenden Schritte umfasst:
    Anordnen einer hochschmelzenden Kernanordnung (30'), die ein erstes Element (32') und ein zweites Element (34'), das mit dem ersten Element verbunden ist, umfasst, in einer Form,
    wobei das erste Element (32') einen in Längsrichtung verlaufenden Schlitz (36') aufweist und das zweite Element (34') einen in Längsrichtung verlaufenden Schlitz (38') aufweist, der einen Abschnitt des ersten Elements (32') aufnimmt, so dass das erste Element (32') und das zweite Element (34') ineinandergreifen, wobei die hochschmelzende Kernanordnung (30') dazu konfiguriert ist, dass sie einen Dichtungsschlitz (16') in einer Turbinenschaufelummantelung herstellt;
    Einkapseln der hochschmelzenden Kernanordnung (30') in einem Wachsausschmelzmodell, das die Form der Turbinentriebwerkskomponente aufweist;
    Herstellen einer keramischen Maskenform um das Wachsausschmelzmodell;
    Entfernen des Wachsausschmelzmodells; und
    Gießen von geschmolzenem Material in die keramische Maskenform, um die Turbinentriebwerkskomponente herzustellen;
    dadurch gekennzeichnet, dass:
    das zweite Element (34') einen planaren Abschnitt (134') und einen nichtplanaren Abschnitt (42') aufweist, der einen gekrümmten Abschnitt angrenzend an den planaren Abschnitt (134') umfasst.
  8. Verfahren nach Anspruch 7, ferner umfassend das Entfernen der hochschmelzenden Kernanordnung (30'), nachdem das geschmolzene Material erstarrt ist, um einen Schlitz in einer Wand eines Abschnitts der Turbinentriebwerkskomponente herzustellen, zum Beispiel mittels eines Säureauslagungsverfahrens.
  9. Verfahren nach Anspruch 7 oder 8, wobei der Schritt des Anordnens das Anordnen einer Vielzahl von hochschmelzenden Kernanordnungen (30') in der Form umfasst.
  10. Verfahren nach einem der Ansprüche 7 bis 9, wobei der Schritt des Anordnens das Anordnen der hochschmelzenden Kernanordnung (30') in einem Abschnitt der Form, die zu verwenden ist, um mindestens eines aus einem äußeren Ummantelungsring und einem inneren Ummantelungsring herzustellen, umfasst.
  11. Verfahren nach einem der Ansprüche 7 bis 10, wobei der Schritt des Gießens eines geschmolzenen Materials das Gießen einer Superlegierung auf Nickelbasis umfasst.
EP11151690.2A 2010-01-25 2011-01-21 Gegossene Ummantelungsschlitze mit Vorwirbelleck Active EP2354464B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US12/692,718 US20110182726A1 (en) 2010-01-25 2010-01-25 As-cast shroud slots with pre-swirled leakage

Publications (3)

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EP2354464A2 EP2354464A2 (de) 2011-08-10
EP2354464A3 EP2354464A3 (de) 2015-01-14
EP2354464B1 true EP2354464B1 (de) 2020-04-01

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Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10450897B2 (en) * 2016-07-18 2019-10-22 General Electric Company Shroud for a gas turbine engine
CN111936725B (zh) * 2018-03-30 2022-08-16 西门子能源全球两合公司 涡轮机护罩段之间的密封布置
US11428160B2 (en) 2020-12-31 2022-08-30 General Electric Company Gas turbine engine with interdigitated turbine and gear assembly

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EP2354464A2 (de) 2011-08-10
EP2354464A3 (de) 2015-01-14
US20110182726A1 (en) 2011-07-28

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