Cast shroud slots with pre-swirled leakage

Description

BACKGROUND

[0001] 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.

[0002] U.S. patent application publication no. US 2008/0145226 A1 and European patent application no. 1935531 A2 both address the use of refractory metal cores to cast slots into a turbine vane shroud, and avoid later manufacturing operations such as electro-discharge machining to cut the shrouds.

[0003] It is desirable to manage turbine shroud leakage flows in ways that minimize leakage mixing and acceleration to rotational speed by injecting the leakage flow at a sharp angle to the shroud to accelerate it to the rotating flow next to the shroud. It is physically difficult to angle the segmentation line of the shroud, without interfering with the structural attachment of the airfoil to the shroud wall. It is desirable to have a variable shroud segmentation cut, that addresses the need for segmentation, and acceleration of the leakage flow associated with the segmentation cut.

[0004] JP S60 22002A discloses a blade structure for preventing secondary flow.

[0005] 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.

SUMMARY

[0006] In accordance with the present disclosure, there is provided 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.

[0007] Still further in accordance with the instant disclosure there is provided 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.

[0008] Other details of the as-cast shroud slots with pre-swirled leakage are set forth in the following detailed description wherein like reference numerals depict like elements.

BRIEF DESCRIPTION OF THE DRAWINGS

[0009] 

Fig. 1 is a top view of a portion of a turbine engine component having a segmented shroud;

Fig. 2 is a top view of a portion of a turbine engine component having a segmented shroud which provides a pre-swirled leakage;

Fig. 3 illustrates a refractory metal core system for forming the segmented shroud of Fig. 2;

Fig. 4 is a top view of the system of Fig. 3;

Fig. 5 is a side view of the system of Fig. 3;

Fig. 6 illustrates an as-cast turbine engine component having a segmented shroud;

Fig. 7 is a sectional view taken along lines 7 - 7 of Fig. 6;

Fig. 8 is a rear view of the as-cast turbine engine component of Fig. 6; and

Fig. 9 is a flow chart illustrating the process for forming the turbine engine component.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT(S)

[0010] 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.

[0011] Fig. 2 illustrates a turbine engine component 10' fabricated in accordance with the present invention. As can be seen from Fig. 2, the turbine engine component 10' has a segmented shroud ring 14' with a plurality of airfoils 12'. However, in this case, 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'.

[0012] 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.

[0013] 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. As can be seen from Fig. 3, 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'. Still further, 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.

[0014] 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.

[0015] 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.

[0016] 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. For example, the surfaces 130' and 132' may be perpendicular to the surfaces 136' and 138'.

[0017] 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. For example, as set forth in Fig. 9, in a step 202, 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. In step 204, the refractory core assembly is encapsulated in a wax pattern having the form of the turbine engine component 10'. In step 206, a ceramic shell mold is formed about the wax pattern. In step 208, the wax pattern is removed. In step 210, molten material is poured into the ceramic shell mold to form said turbine engine component. In step 212, the airfoils and the turbine engine component may be removed from the ceramic shell mold. In 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.

[0018] In step 202, 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.

[0019] If desired, in step 202, 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.

[0020] If desired, in step 210, 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.

[0021] The resultant turbine engine component 10' formed by the foregoing process, as shown in Figure 6, 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. As can be seen from Fig. 6, 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'.

[0022] Fig. 8 illustrates the segmented shroud ring 14' with a flow of fluid 54' being discharged from the slot 16'. As can be seen from this figure, the fluid is pre-swirled and flows in the direction of rotation. Such a flow reduces losses.

[0023] 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.

[0024] Using the 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.

[0025] 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.

[0026] It is apparent that there has been described herein as-cast shroud slots with pre-swirled leakage. While the disclosure has been set out in the form of specific embodiments, other unforeseeable variations, modifications, and alternatives may become apparent to those skilled in the art having read the foregoing specification. Accordingly, it is intended to embrace those unforseseen alternatives, modifications, and variations as fall within the broad scope of the appended claims.

Claims

1. A system (30') of elements configured to interlock as a refractory metal core assembly configurable in a die for forming a seal slot (16') having a curved trailing edge portion in a turbine vane shroud assembly, said system comprising a first element (32') having a longitudinally extending slot (36') and a second element (34') having a longitudinally extending slot (38') for receiving a portion of said first element (32') so that the first element (32') and the second element (34') interlock, characterised in that said second element (34') has a planar portion (134') and a non-planar portion (42') comprising a curved portion adjacent said planar portion (134'), wherein said curved portion (42') is configured in a shape to form a curved trailing edge portion of a said seal slot (16').

2. The system of claim 1, wherein said first element (32') has a pair of opposed surfaces and said second element (34') has a pair of opposed surfaces which are oriented at an angle with respect to said surfaces of said first element (32').

3. The system of claim 1 or 2, wherein said curved portion (42') is configured to form a trailing edge portion of a said seal slot (16') in a shape to direct leakage flow discharged from said seal slot (16') in a gas turbine engine segmented vane shroud ring to be more aligned with a gas turbine engine rotating flow next to the turbine vane shroud, compared with a flow entering said seal slot (16').

4. The system of any preceding claim, wherein said first element (32') has a first length and said second element (34') has a second length greater than said first length.

5. The system of any preceding claim, wherein each of said first and second elements (32', 34') is formed from a refractory metal material.

6. The system of any preceding claim, wherein said second element (34') has a portion (144') with a twist.

7. A process for forming a turbine engine component comprising a turbine vane shroud, the process comprising the steps of:

placing a refractory core assembly (30') comprising a first element (32') and a second element (34') joined to said first element in a die, the first element (32') having a longitudinally extending slot (36') and the second element (34') having a longitudinally extending slot (38') receiving a portion of the first element (32') so that the first element (32') and the second element (34') interlock, the refractory core assembly (30') being configured to form a seal slot (16') in a said turbine vane shroud;

encapsulating said refractory core assembly (30') 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 (34') has a planar portion (134') and a non-planar portion (42') comprising a curved portion adjacent said planar portion (134').

8. The process of claim 7, further comprising removing said refractory core assembly (30') after said molten material has solidified so as to from a slot in a wall of a portion of said turbine engine component, for example using an acid leach operation.

9. The process of claim 7, or 8 wherein said placing step comprises placing a plurality of refractory core assemblies (30') in said die.

10. The process of any of claims 7 to 9, wherein said placing step comprises placing said refractory core assembly (30') in a portion of said die to be used to form at least one of an outer shroud ring and an inner shroud ring.

11. The process of any of claims 7 to 10, wherein said molten material pouring step comprises pouring a nickel based superalloy.

Ansprüche

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.

Revendications

1. Système (30') d'éléments configurés pour se verrouiller en tant qu'ensemble noyau métallique réfractaire configurable dans une matrice pour former une encoche de joint (16') ayant une partie de bord de fuite incurvée dans un ensemble protection brute d'aube de turbine, ledit système comprenant un premier élément (32') ayant une encoche s'étendant longitudinalement (36') et un second élément (34') ayant une encoche s'étendant longitudinalement (38') pour recevoir une partie dudit premier élément (32') de sorte que le premier élément (32') et le second élément (34') se verrouillent, caractérisé en ce que ledit second élément (34') a une partie plane (134') et une partie non plane (42') comprenant une partie incurvée adjacente à ladite partie plane (134'), dans lequel ladite partie incurvée (42') est configurée en une forme pour former une partie de bord de fuite incurvée d'une dite encoche de joint (16').

2. Système selon la revendication 1, dans lequel ledit premier élément (32') a une paire de surfaces opposées et ledit second élément (34') a une paire de surfaces opposées qui sont orientées selon un angle par rapport auxdites surfaces dudit premier élément (32').

3. Système selon la revendication 1 ou 2, dans lequel ladite partie incurvée (42') est configurée pour former une partie de bord de fuite d'une dite encoche de joint (16') en une forme pour diriger un écoulement de fuite évacué de ladite encoche de joint (16') dans un anneau de protection brute d'aube segmenté de moteur à turbine à gaz pour être plus aligné avec un écoulement tournant de moteur à turbine à gaz à côté de la protection brute d'aube de turbine, par rapport à un écoulement entrant dans ladite encoche de joint (16').

4. Système selon une quelconque revendication précédente, dans lequel ledit premier élément (32') a une première longueur et ledit second élément (34') a une seconde longueur supérieure à ladite première longueur.

5. Système selon une quelconque revendication précédente, dans lequel chacun desdits premier et second éléments (32', 34') est formé à partir d'un matériau métallique réfractaire.

6. Système selon une quelconque revendication précédente, dans lequel ledit second élément (34') a une partie (144') avec une torsion.

7. Procédé de formation d'un composant de moteur à turbine comprenant une protection brute d'aube de turbine, le procédé comprenant les étapes :

de placement d'un ensemble noyau réfractaire (30') comprenant un premier élément (32') et un second élément (34') relié audit premier élément dans une matrice, le premier élément (32') ayant une encoche s'étendant longitudinalement (36') et le second élément (34') ayant une encoche s'étendant longitudinalement (38') recevant une partie du premier élément (32') de sorte que le premier élément (32') et le second élément (34') se verrouillent, l'ensemble noyau réfractaire (30') étant configuré pour former une encoche de joint (16') dans une dite protection brute d'aube de turbine ;

l'encapsulation dudit ensemble noyau réfractaire (30') dans un motif de cire ayant la forme dudit composant de moteur à turbine ;

la formation d'un moule à coque céramique autour dudit motif de cire ;

le retrait dudit motif de cire ; et

le versement du matériau fondu dans ledit moule à coque céramique pour former ledit composant de moteur à turbine ; caractérisé en ce que :

ledit second élément (34') a une partie plane (134') et une partie non plane (42') comprenant une partie incurvée adjacente à ladite partie plane (134').

8. Procédé selon la revendication 7, comprenant en outre le retrait dudit ensemble noyau réfractaire (30') après que ledit matériau fondu s'est solidifié de manière à former une encoche dans une paroi d'une partie dudit composant de moteur à turbine, par exemple à l'aide d'une opération de lixiviation acide.

9. Procédé selon la revendication 7 ou 8, dans lequel ladite étape de placement comprend le placement d'une pluralité d'ensembles noyau réfractaire (30') dans ladite matrice.

10. Procédé selon l'une quelconque des revendications 7 à 9, dans lequel ladite étape de placement comprend le placement dudit ensemble noyau réfractaire (30') dans une partie de ladite matrice à utiliser pour former au moins l'un d'un anneau de protection brute extérieur et d'un anneau de protection brute intérieur.

11. Procédé selon l'une quelconque des revendications 7 à 10, dans lequel ladite étape de versement de matériau fondu comprend le versement d'un superalliage à base de nickel.

Drawing