Patent Application
20150267557
This application claims priority to European application 14160697.0 filed Mar. 19, 2014, the contents of which are hereby incorporated in its entirety.
The present invention relates to the field of rotor blades or guide vanes of a turbo-machine, especially of a gas or steam turbine. The final aim of the present invention is providing adequate cooling in a rotor blade or guide vane airfoil improving the cooling flow control and enabling insert fits.
Accordingly, the present invention relates to a rotor blade or guide vane airfoil assembling of a gas or steam turbine and refers fundamentally to a specific or modular arrangement of airfoil inserts within the cavity of the respective airfoil portion.
Basically, the specific or modular arrangement of airfoil inserts within the cavity of the respective airfoil portion consisting of replaceable and/or non-replaceable inserts. Besides the used airfoil inserts, the rotor blade or guide vane comprising additionally substitutable and non-substitutable flow-applied and no flow-applied elements.
Accordingly, the present invention relates to a turbine blade, namely as rotor blade or guide vane, with a hollow airfoil portion having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, a convex suction side, a concave pressure side. At least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion. The insert further comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto.
U.S. Pat. No. 8,182,203 B2 discloses a turbine blade including an airfoil; a supply channel extending through the interior of the airfoil in the span direction, through which cooling fluid flows; a pin fin channel extending from the supply channel along the center line of the airfoil toward the trailing edge of the airfoil and opening at the trailing edge to the exterior of the airfoil; a plurality of gap pin fins projecting from a pair of opposing inner walls that constitute the pin fin channel at a region at the supply channel side of the pin fin channel and forming a gap there between extending in the span direction; pin fins connecting the pair of opposing inner walls at a region at the trailing edge side of the pin fin channel; and an insertion portion disposed in the gap to decrease the area of the channel of the cooling fluid at the region at the supply channel side of the pin fin channel.
With the turbine blade and the gas turbine of the present disclosure, the insertion portion is disposed in the gap formed between the gap pin fins. Therefore, the cross-sectional area of the channel at the supply channel side of the pin fin channel, through which cooling fluid flows, decreases as compared with a case in which the insertion portion is not disposed, so that the velocity of the cooling fluid at the region at the supply channel side increases. This increases the cooling efficiency at the region at the supply channel side, which improves the cooling efficiency of the pin fin channel, thus improving the cooling performance of the turbine blade.
Referring to EP 2 492 442 A2 a vane is provided for directing hot gases in a gas turbine engine. The vane includes a hollow aerofoil portion, which in use spans the working gas annulus of the engine. The vane further includes an impingement tube which forms a covering over the interior surface of the aerofoil portion and which has jet-forming apertures formed therein for the production of impingement cooling jets. The impingement tube includes two tube portions which are separately insertable into position into the aerofoil portion to form the covering. The impingement tube further includes an expansion member which, when the tube portions are in position in the aerofoil portion, is locatable in the aerofoil portion to urge each tube portion outwardly and thereby holds the tube portions in position against the aerofoil portion.
Referring to U.S. Pat. No. 8,231,329 B2 a turbine blade with a generally hollow airfoil having an outer wall that defines a chamber for receiving cooling air, the airfoil comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, a convex suction side, a concave pressure side, and an insert disposed within the chamber that is configured to initially receive at least a portion of the cooling air entering the chamber and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall, the insert further comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto, wherein the chamber and insert narrow as they extend toward the trailing edge, the insert eventually terminating and the chamber eventually terminating at a pin array section; wherein a first distance exists that comprises the generally axial distance between the position of downstream termination point of the insert and the position of an upstream beginning point of the pin array section, wherein the pin array section, at a downstream end, comprises a plurality of openings that define an inlet to a plurality of trailing edge cooling apertures, and wherein the chamber, the insert, and the pin array section are configured such that the first distance is approximately minimized.
U.S. Pat. No. 7,452,182 B2 relates to a modular guide vane assembly. The vane assembly includes an airfoil portion, an outer platform and an inner platform. The airfoil portion can be made of at least two segments. Preferably, the components are connected together so as to permit assembly and disassembly of the vane. Thus, in the event of damage to the vane, repair involves the replacement of only the damaged subcomponents as opposed to the entire vane. The modular design facilitates the use of various materials in the vane, including materials that are dissimilar. Thus, suitable materials can be selected to optimize component life, cooling air usage, aerodynamic performance, and cost. Because the vane is an assemblage of smaller sub-components as opposed to one unitary structure, the individual components of the vane can be more easily manufactured and more intricate features can be included. According to this document, one end of the airfoil can be received within a recess in one of the inner and outer platforms. The assembly can further include a seal provided between the recesses and at least one of the radial ending of the airfoil and the outer peripheral surface of the airfoil proximate the radial end. As a result, hot gas infiltration or cooling air leakage can be minimized. In such case, one or more of the airfoil segments, the inner platform and/or the outer platform can be made of Intermetallics, Oxide Dispersion Strengthened (ODS) alloys, single-crystal metals, advanced Superalloys, metal matrix composites, ceramics or CMC.
The inventive idea of the present invention leaves the use of typical rotor blade or guide vanes assembling consisting of an airfoil portion, an inner and an outer platform, also called shroud, made in one piece as depicted and explained in connection with notorious state of the art.
Especially, by using a rotor blade or guide vane which can be assembled by at least two separate parts, i.e. a separate airfoil portion and outer platform and a separate inner platform, on the one hand preconditions are created to provide interchange ability or repairing and/or reconditioning of the identified separate parts, modules, elements without replacing the whole rotor blade or guide vane.
On the other hand, it is also possible to use rotor blades or guide vanes of three separable parts, i.e. outer platform, airfoil portion and inner platform. In a separate process the various parts or modules or elements of the guide vane may be repaired and/or reconditioned.
Additionally, the present invention describes an improved rotor blade or guide vane assembling of a gas or steam turbine on the basis of a modular structure comprising fundamentally an airfoil portion, inner platform, outer platform, whereas the airfoil portion and/or the platforms having at its one end means for the purpose of an interchangeable connection of rotor blade or vane guide elements, whereas the connection of rotor blade or guide vane elements having a permanent or semi-permanent fixation with respect to the airfoil portion in radially or quasi-radially extension and with respect to the axis of the gas or steam turbine, whereas the assembling of the airfoil portion in connection with platforms based on a friction-locked bonding actuated by adherence interconnecting, or the assembling of the airfoil portion in connection with platforms based on the use of a metallic and/or ceramic surface the fixing guide vane elements to each other, or the assembling of the airfoil portion in connection with platforms based on force closure means with a detachable or permanent connection, whereas at least the airfoil portion comprising at least one outer hot gas path liner encasing at least one part of the airfoil portion.
Moreover and basically, the present invention uses same or similar assemblies to determine the various possible connection of various configured airfoil inserts within the cavity of the airfoil portion. In order to decrease the size of the trailing edge channel inlet, at the end of the respective airfoil insert, one or more additional airfoil insert(s) can be used.
In this context, the additional airfoil insert(s) could be inserted and slide in the trailing edge region before to put in place a main airfoil insert. The additional airfoil insert(s) could optionally be cast in. The insertion of the airfoil additional insert(s) is performed by a cascade principle with respect to their size, namely:
The additional insert (see
At least one main airfoil insert may be inserted afterwards. It is also possible to proceed conversely.
Accordingly, at least one main insert comprising at least one additional insert which is inserted from the outside and transferred into the cavity at an intermediate position, and then moved in direction to the trailing edge and fixed to predetermined position, wherein the additional insert forms the size of the trailing edge channel inlet at the end of the main insert.
Moreover, at least one main insert comprising at least one additional insert which forms the size of the trailing edge channel inlet at the end of the main insert. This additional insert consists of a structured unitary body.
Different sized inserts can be arranged in the transverse direction of rotor blade or guide vane.
Various gaps between the airfoil inserts can be provided in all directions within the airfoil cavity on a case by case basis.
A joining assembling referring to airfoil inserts can be mechanically secured, or the joining assembling can use a shrinking process.
Fundamentally, the detachable or permanent connection comprising a force closure with bolt or rivet, or by HT brazing, active brazing, soldering. Additionally, an individual insert can be made of one piece or of a composite structure.
Furthermore, the inserts are able to resist the caloric and physical stresses, wherein the mentioned means are holistically or on their part interchangeable among one another.
Accordingly, one of the basic idea of the invention consists to split one or more inserts within the cavity of the airfoil portion in multiple inserts in order to better adapt at the rotor blade or guide vane geometries, regardless of whether the respective rotor blade or guide vane consist of a unique body or a modular structure.
In this context, the invention provides adequate cooling in the airfoil, improving the flow control and enabling the insert fits.
Having a multiple airfoil inserts configuration as the one proposed embodiment in this invention disclosure would allow improving the design flexibility and the part performance.
In order to decrease the size of the trailing edge channel inlet at the end of the insert one or more additional inserts can be used. The additional insert(s) could be inserted and slide in the trailing edge region before to put in place the main insert(s). The additional insert(s) could optionally be cast in.
In one embodiment of the present invention, the inserts can be made of the same material as the respective airfoil portion in which they are intercalated, such as IN939 alloy and ECY768 alloy. The inserts can be made of a material that may or may not have a greater resistance to heat compared to the material of the airfoil portion. For example, the inserts can be made of a material with a lower heat resistance than the material of the receiving airfoil portion. The inserts can be made from an inexpensive material so that the cost of a replacement insert would not significantly add to the overall costs over the life of the engine.
For insertion or removal purpose of the airfoil portion inserts it is possible to handle the mentioned airfoil portion inserts only at its radially outwards directed end which is a remarkable feature for performing maintenance work at the turbine stage.
The term “radial,” as used herein, is intended to mean radial to the turbine when the rotor blade or guide vane assembling is installed in its operational position.
Furthermore, a manner of attaching the airfoil portion and their insert portions to the inner respectively outer platform consists in doing the fact, that the radially end of the mentioned element can be received in a recess provided in the respective platform. The mentioned recesses can be substantially airfoil-shaped so as to correspond to the outer contour of the airfoil portion and airfoil inserts. Thus, the airfoil portion assembly, including optionally an outer shell arrangement, can be trapped between the inner platform and the outer platform.
One of the most important solutions of the invention is to provide at least one outer shell and, if necessary and needed and according to individual operative requirements or different operating regimes, at least one no flow-applied intermediate shell in connection with the airfoil inserts for modular variants of the original airfoil portion. Function of the airfoil carrier is to carry mechanical load from the airfoil module. In order to protect the airfoil carrier with respect to the high temperature and separate thermal deformation from the airfoil module, an outer and, additionally, an intermediate hot gas path shells are introduced.
If several superimposed shells with respect to the airfoil portion or their inserts are provided, they can be built with or without intermediate spaces between each other.
The mentioned shells can be made of at least two segments. Preferably, the components forming the shell are connected together so as to permit assembly and disassembly of shell, shell components, airfoil portion and airfoil inserts of rotor blade or guide vane.
If the airfoil portion and airfoil inserts are internally cooled with a cooling medium at a higher pressure than the hot combustion gases, excessive cooling medium leakage into the hot gas path can occur. To minimize such concerns, one or more additional seals can be provided in connection with the shell arrangement. The seals can be at least one of rope seals, W-shaped seals, C-shaped seals, E-shaped seals, a flat plate, and labyrinth seals. The seals can be made of various materials including, for example, metals and ceramics.
The main advantages of the present invention are as follows:
The above explained statements together with the other aspects of the present disclosure, along with the various features that characterize the present invention, are pointed out with particularity in the present disclosure. For a better understanding of the present disclosure, its operating advantages, and its uses, reference should be made to the accompanying drawings and descriptive matter in which there are illustrated exemplary embodiments of the present disclosure.
The advantages and features of the present disclosure will be better understood with reference to the following detailed description and claims taken in conjunction with the accompanying drawing, wherein like elements are identified with like symbols, and in which:
As shown in
The airfoil portion 100 has an integral cavity 101, which is a hollow formed at the leading edge 102 and extending in the flow direction of the airfoil portion 100 to the trailing edge 103. At least in the region of the leading edge 102 the external wall 104 of the airfoil portion 100 comprising a number of film-cooling holes 105 communicating with the front cavity 101. In other words, the airfoil portion 100 has, in its interior, a first integrally cavity 101 extending in the flow direction or the airfoil portion 100. The inner cavity 101 can be provided with at least one partition (not shown) in the manner that the partition may be divided the hollow portion into a front cavity and a rear cavity.
A cooling fluid derived from the exterior, for example compressed air extracted from the compressor, cools adequately the structure of the airfoil portion 100.
In the cavity 101 a main hollow (205) insert 106 is disposed at a predetermined space from the inner wall of the cavity 101. On the other hand, if the cavity is provided with partitions, in the rear cavity space a rear insert is also disposed at a predetermined space from the inner wall of the rear cavity.
As shown in
Furthermore, the film cooling holes 105 are formed from the front cavity 101 to the exterior as slanting holes inclined from the leading edge 102 to the trailing edge 103
Moreover, the rear cavity of the airfoil portion 100 is provided with a pin fin channel 109, which is a hollow extending from the rear cavity 101 toward the trailing edge 103 along a center line of the airfoil 100 (not shown) and which is a region in which gap pin fins 110 and pin fins 111 are provided.
The gap pin fins 110 are a plurality of substantially columnar members protruding from regions at the rear cavity side of the pin fin channel 109, the regions being a pair of inner walls constituting the pin fin channel 109. The amount of protrusion of the gap pin fins 110 from the above-described inner walls is set so as to form a gap between the gap pin fins 110 into which the end portion of the rear or additional insert 200 can be inserted.
The pin fins 111 are a plurality of substantially columnar members that connect regions at the trailing edge 103 side of the pin fin channel 109, the regions being the pair of inner walls constituting the pin fin channel 109. The shape and arrangement of the pin fins 111 can be known ones and are not particularly limited.
The pin fin channel 109 is a channel in the rear cavity in the region of the trailing edge 103, through which cooling fluid flows after being used for impinging cooling, and constitutes a structure related to pin fin cooling for cooling the vicinity of the trailing edge 103 of the airfoil portion 100 and opens to the exterior at the trailing edge 103.
As shown in
The front of the main insert 106 constitutes a structure related to impinging cooling for cooling the leading edge 102 and the other inner wall of the airfoil portion 100, together with the front and the subsequent cavity 101. The front of the main insert 106 consists of a substantially cylindrical member having a cross-sectional form similar to the cross-sectional form of the front cavity 101. Furthermore, the front of the main insert 106 has a plurality of discharge holes 113 through which the cooling fluid flowing there through spouts against the inner wall of the front cavity 101
If the airfoil portion 100 is provided with partitions the rear part of the insert constitutes also a structure related to impinging cooling, like the front insert, for cooling the respective side of the airfoil portion 100. The rear insert consists also of a substantially cylindrical member having a cross-sectional form similar to the cross-sectional form of the rear part of the cavity.
In order to decrease the size of the trailing edge channel inlet 109, at the end of the respective airfoil main insert 106, an additional airfoil insert 200 is used.
One possibility consists in the fact that the additional airfoil insert 200 is inserted and slide in the trailing edge region 103 before to put in place the main airfoil insert 106. The additional airfoil insert 200 could optionally be cast in. The insertion of the airfoil additional insert 200 is performed by a cascade principle with respect to their size, namely:
The additional airfoil insert 200 is inserted from the outside 201 and transferred 201a into the cavity 101 at an intermediate position 202, and then moved 202a in direction to the trailing edge 103 and fixed to predetermined position 203.
The main airfoil portion insert 106 may be inserted afterwards. But it is also possible to proceed conversely. In the last mentioned case, the additional insert 200 is provided with a transversally elasticity 204, so that it can be pushed over the end side constriction of the main insert 106 until it reaches its final position 203. The connection between the main 106 and the additional insert 200 is designed accordingly, even in the case in which the additional insert 200 does not have any transversally elasticity 204. Thus, the connection can be obtained mechanically, for example with introduction of fixing members (not shown), positioned in the region of the both inserts.
Furthermore, the additional insert 200 which is inserted from the outside 201 and transferred 201a into the cavity 101 at an intermediate position, and then it can be moved alternatively in direction to the leading edge 102 and fixed to final predetermined position. Additionally, the additional insert 200 can be inserted from the underside of the airfoil portion or is an element of the cavity of the airfoil portion, and then it can be moved in the direction of the trailing edge or leading edge and fixed to final predetermined position. Accordingly, the additional insert 200 forms the size of the trailing edge cavity at the end of the main insert 106, or the additional insert forms the leading edge cavity between inner wall of the airfoil portion and subsequent main insert.
Summary, an airfoil portion 100 of a rotor blade or guide vane of a turbo-machinery having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, suction side, a pressure side. At least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion. Furthermore, the insert comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto. A portion of the cooling air exits the airfoil portion through a plurality of film cooling apertures formed through the outer wall and/or a portion of the cooling fluid exits the airfoil at the trailing edge. At least one main insert 106 comprising at least one additional insert 200 which is inserted as a first option from the outside 201 and transferred 201a into the cavity 101 at an intermediate position 202, and then the additional insert is moved 202a in direction to the trailing edge 103 or leading edge 102 and fixed to final predetermined position 203. A second option consists in the fact that the additional insert 200 can be inserted from the underside of the airfoil portion or consists of an element of the cavity of the airfoil portion. Accordingly, the additional insert is moved in the direction of the trailing edge or leading edge and fixed to final predetermined position, wherein the additional insert 200 forms at least one size of the trailing edge cavity at the end of the main insert 106, or the additional insert forms at least one leading edge cavity between inner wall of the airfoil portion and subsequent disposed main insert.
As shown in
Summary, an airfoil portion 100a of a rotor blade or guide vane of a turbo-machinery having an outer wall that defines a cavity for receiving cooling air, the airfoil portion comprising a leading edge that resides in an upstream direction, a trailing edge that resides in a downstream direction, a convex suction side, a concave pressure side, and at least one insert disposed within the cavity that is configured to initially receive at least a portion of the cooling air entering the chamber of the insert and direct the cooling air through a plurality of insert apertures to cool the inner surface of the outer wall of the airfoil portion. The insert further comprising a configuration that generally conforms to the contour of the outer wall of the chamber but in spaced relation thereto. Additionally, a portion of the cooling air exits the airfoil portion through a plurality of film cooling apertures formed through the outer wall and/or a portion of the cooling fluid exits the airfoil at the trailing edge. At least one main insert 106a comprising at least one additional insert 250 which forms the size of the trailing edge cavity at the end of the main insert 106a and/or at least one main insert comprising at least one additional insert which form the size of the leading edge cavity at the initiation of the main insert.
As shown in
With respect to embodiments according to
Summary, an airfoil portion 100b of a rotor blade or guide vane of a turbo-machinery having an outer wall 104 which defines the cavity (see also
The main and/or additional inserts extend to radially or quasi-radially and/or transversally or quasi-transversally direction of the airfoil portion and are sectioned and having different shapes or profiles along one or more orientations of the airfoil portion, see
The airfoil portion having in radially or quasi-radially direction compared to the axis of the turbo-machinery a pronounced or swirled or tailored aero-dynamic profile. (see
While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, preferred or advantageously in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.
| Number | Date | Country | Kind |
|---|---|---|---|
| 14160697.0 | Mar 2014 | EP | regional |
