The present invention relates to a rotor blade or guide vane assembly of a turbomachine, especially of a gas turbine engine. In particular, the invention relates to a modularly designed blade or vane of a turbomachine with one or more interchangeable elements or modules. In addition, the invention is applicable for other turbomachine parts, like heat shields, inner platform, outer platform, airfoil carrier, etc.
Additionally, the invention relates to manufacturing, reconditioning, joining and disassembling methods with respect to turbomachine components.
EP 2 204 537 A2 discloses a guide vane for a gas turbine. This vane consists of a composite airfoil structure. The composite airfoil structure may have an opening. The vane may include a spar. The spar may have a body, which may be disposed within the opening. A standoff structure may be disposed within the opening. In some nonlimiting embodiments, a cooling air gap may be defined between the body and an internal surface of the composite airfoil structure.
U.S. Pat. No. 8,235,670 B2 shows a ceramic matrix composite (CMC) airfoil, assembled from a pressure side wall and a suction side wall, joined by interlocking joints at the leading edge and trailing edge of the airfoil to form a tapered thin trailing edge. The trailing edge is thinner than a combined thickness of the airfoil walls. One or both of the interlocking joints may be formed to allow only a single direction of assembly, as exemplified by a dovetail joint. Each joint includes keys on one side and respective keyways on the other side. Each keyway may have a ramp that eliminates indents in the airfoil outer surface that would otherwise result from the joint.
According to US 2011/0142684 A1 a rotor blade airfoil is formed by a first process using a first material. A platform is formed by a second process using a second material that may be different from the first material. The platform is assembled around a shank of the airfoil. One or more pins extend from the platform into holes in the shank. The platform may be formed in two portions and placed around the shank, enclosing it. The two platform portions may be bonded to each other. Alternatively, the platform may be cast around the shank using a metal alloy with better castability than that of the blade or shank, which may be specialized for thermal tolerance. The pins bear load from the under section of the airfoil.
According to US 2011/0142639 A1 a turbine airfoil extends from a shank. A platform brackets or surrounds a first portion of the shank Opposed teeth extend laterally from the platform to engage into respective slots in a disk. Opposed teeth extend laterally from a second portion of the shank that extends below the platform to engage other slots in the disk. Thus the platform and the shank independently support their own centrifugal loads via their respective teeth. The platform may be formed in two portions that are bonded to each other at matching end-walls and/or via pins passing through the shank. Coolant channels may pass through the shank beside the pins.
EP 2 189 626 B1 describes to a rotor blade arrangement, especially for a gas turbine, which rotor blade arrangement can be fastened on a blade carrier and comprises in each case a blade airfoil element and a platform element, wherein the platform elements of a blade row form a continuous inner shroud. With such a blade arrangement, a mechanical decoupling, which extends the service life, is achieved by blade airfoil element and platform element being formed as separate elements and by being able to be fastened in each case separately on the blade carrier.
US 2011/268582 A1 relates to a blade, comprising a blade airfoil which extends in the longitudinal direction of the blade along a longitudinal axis. The blade airfoil, which is delimited by a leading edge and a trailing edge in the flow direction, merges into a shank at the lower end beneath a platform which forms the inner wall of the hot gas passage, and ends in a customary blade root portion with a fir-tree-shaped cross-sectional profile by which the blade can be fastened on a blade carrier, especially on a rotor disk, by inserting into a corresponding axial slot (see, for example, FIG. 1 of U.S. Pat. No. 4,940,388).
Cooling passages extend inside the blade airfoil for cooling the blade with a cooling medium, particularly cooling air, via a feed hole which is arranged on the shank. The shank, similar to the blade airfoil, has a concave and a convex side. The feed hole, which extends obliquely upwards into the interior of the blade airfoil, opens into the outside space on the convex side of the shank. As shown in
In US 2013/0089431 A1 a blade airfoil for a turbine system is disclosed. The blade airfoil includes a first body having exterior surfaces defining a first portion of an aerodynamic contour of the blade airfoil and formed from a first material. The blade airfoil further includes a second body having exterior surfaces defining a second portion of an aerodynamic contour of the blade airfoil, the second body is coupled to the first body and formed from a second material having a different temperature stability than the first material. In another embodiment, a nozzle for a turbine section of a turbine system is disclosed. The nozzle includes a blade airfoil having exterior surfaces defining an aerodynamic contour, the aerodynamic contour comprising a pressure side and a suction side extending between a leading edge and a trailing edge. The blade airfoil includes a first body having exterior surfaces defining a first portion of the aerodynamic contour of the blade airfoil and formed from a first material. The blade airfoil further includes a second body having exterior surfaces defining a second portion of the aerodynamic contour of the blade airfoil, the second body is coupled to the first body and formed from a second material having a different temperature stability than the first material.
U.S. Pat. No. 5,700,131 shows an internally cooled turbine blade for a gas turbine engine that is modified at the leading and trailing edges to include a dynamic cooling air radial passageway with an inlet at the root portion and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the blade airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film-hole coverage throughout the radial length. Trip strips may be employed to augment the pressure drop distribution. Also well known is that the engine's efficiency increases as the pressure ratio of the turbine increases and the weight of the turbine decreases. Increasing the speed of the turbine also increases the blade airfoil loading and, of course, satisfactory operation of the turbine is to stay within given blade airfoil loadings. The blade airfoil loadings are governed by the cross sectional area of the turbine multiplied by the velocity of the tip of the turbine squared. Obviously, the rotational speed of the turbine has a significant impact on the loadings. The spar/shell construction contemplated by this invention affords the turbine engine designer the option of reducing the amount of cooling air that is required in any given engine design. And in addition, allowing the designer to fabricate the shell from exotic high temperature materials that heretofore could not be cast or forged to define the surface profile of the blade airfoil section. In other words, by virtue of this invention, the shell can be made from Niobium or Molybdenum or their alloys. In addition, because of the efficacious cooling scheme of this invention, the shell portion could be made from ceramics or more conventional materials and still present an advantage, because a lesser amount of cooling air would be required.
EP 2 642 076 shows a connecting system for metal components and CMC components, a turbine blade retaining system and rotating component retaining system are provided. The connecting system includes a retaining pin, a metal foam bushing, a first aperture disposed in the metal component, and a second aperture disposed in the ceramic matrix composite component. The first aperture and the second aperture are configured to form a through-hole when the metal component and the ceramic matrix composite component are engaged. The retaining pin and the metal foam bushing are operably arranged within the through-hole to connect the metal component and the ceramic matrix composite component.
U.S. Pat. No. 8,366,398 B1 shows a turbine blade or vane for a gas turbine engine, especially for a large frame industrial gas turbine engine, including a shell secured to a spar through a number of hooks that extend from both the shell and the spar.
The technical problem solved by the present invention relates to joining an outer shell to an inner core structure of an airfoil of a rotor blade or a guide vane of a turbomachine, especially a gas turbine. In addition, the present invention solves the problem of assembling or disassembling such a rotor blade or guide vane.
The inner core structure of the blade or vane includes the following basic features:
The outer shell and, if used, an intermediate shell include the following basic features:
Joining, manufacturing, reconditioning, disassembling processes include the following basic features:
The invention uses e.g. a TBC (thermal barrier coating) layer applied on the metallic surface of the heat exposed component and suggests to join at least one plate-like heat resistant component made of ceramic material onto said TBC-layer which provides a continuous surface of ceramic material to obtain a heat resistant cover on the TBC to reduce the temperature stress of TBC during heat exposure of the heat exposed component. So, the TBC acts as layer or inter-layer between the metallic surface of the heat exposed component and with respect to at least one plate-like heat resistant component.
A directly provided joining in way of brazing of highly porous ceramic plate-like heat resistant component onto metallic substrates could only be achieved in a very limited porosity range of the ceramic material which is characterized by the range between 10 vol % and 90 vol % by using very high brazing temperature over 850° C.
Successful joining by way of brazing of ceramic materials with other porosity ranges onto metal substrates requires a different brazing strategy. According to the invention it is proposed to use a layer of thermal barrier coating (TBC-layer) which has two functions: Firstly, to reduce the thermal mismatch stress to be taken by the ceramic and thereby allow brazing of ceramics with a broader porosity range from 1 to 90 vol %, and, secondly, to add a residual life-time to the heat exposed component in case of a loss of the ceramic plate-like heat resistant component.
The joining of the ceramic plate-like thermally resistant components onto the continuous surface of the layer of thermal barrier coating (TBC-layer) can be performed by brazing, using a reactive air braze or an active braze alloy. It is also possible that the plate-like thermally resistant component is joined on said TBC-layer by a diffusion braze, whereby at least one of the surface of the ceramic material of said TBC-layer or the plate-like thermally resistant component is metallized.
Accordingly, the invention relates to a method of applying or reconditioning metallic surface of thermally exposed components of a cast, machined, forged rotor blade or guide vane of a gas turbine respectively turbomachine, particularly with regard to a flow applied outer shell, airfoil substructure or spar, anchoring members with respect to rotor or stator and additional structured components. The applying or reconditioning process is directed to reduce the temperature stress on the metallic surface and is created by attaching a ceramic insert and/or fulfilled by fitting a coating treatment.
The metallic surface according to such a TBC-layer is directly or indirectly applied, and providing a continuous surface of ceramic material, wherein at least one plate-like thermally resistant component made of ceramic material is joined onto said continuous surface of ceramic material to obtain a thermally resistant cover onto such layer for reducing its temperature stress during the exposure of the thermally exposed component.
The plate-like thermally resistant component is joined on the TBC-layer by a braze, which consists of a reactive air braze or active braze alloy; or the plate-like thermally resistant component is joined on such layer by a diffusion braze and at least one surface of the ceramic material of said layer or the plate-like thermally resistant component is metallized; or the plate-like thermally resistant component is joined on such layer by a cement type adhesive.
Moreover, a multitude of single plate-like thermally resistant components is joined onto the surface of the TBC-layer of such that the plate-like thermally resistant components are arranged side by side covering at least a part of the surface of the layer of thermal barrier coating.
Generally, the plate-like thermally resistant component provides a surface opposed to the layer of thermal barrier coating having a shape of aerodynamic function.
It is further an object of the invention to provide an enhanced method for fixing thermally resistant components, preferably in the form of plate-like ceramic tiles each limited by a peripheral boundary edge, on the surface of a thermally exposed component using a molten solder, which enables that the thermally resistant components are fastened on the surface in a solid and durable way and in particular free from thermal stresses.
According to an advantageous method for fixing a thermally resistant component, like a ceramic tile, on a surface of a thermally exposed component, for example with respect to the outer shell, by means of brazing of at least a part of a surface of the ceramic tile limited by a peripheral boundary edge on the surface of the thermally exposed component using a molten solder is characterized by metallizing the surface of the ceramic tile at least with the exception of an edge area comprising the peripheral boundary edge of the ceramic tile. After metallizing step the metallized surface of the ceramic tile is brazed to the surface of the thermally exposed component in which the peripheral boundary edge of the ceramic tile remains excluded from the braze joint respectively solder joint.
Additionally, it is proposed to work with a braze-metal alloy as solder material which does not wet the ceramic surface of the thermally resistant component unless such surface has undergone metallization before. Thus, the braze area respectively solder area, which corresponds to the joint area, can be defined in shape and size by the metallization process during which a metal layer is coated onto a defined area of the surface of the ceramic tile.
A further alternative for joining the plate-like thermally resistant component onto the continuous surface of the TBC-layer can be realized by a cement type adhesive braze-material.
A single airfoil of a rotor blade or guide vane of turbomachine, e.g. a gas turbine, may be covered completely with a TBC-layer. Additionally, a thermally resistant component made of ceramic material may be attached in the area of the leading edge of the airfoil. Between the ceramic thermally resistant component and the TBC-layer of the airfoil a joint layer which preferably consists of a metallic braze layer, a reactive air braze or an active braze alloy is applied.
The outer shell can be applied to the airfoil inner core structure (spar) in circumferential assembly. The outer shell represents the aero-dynamical profile of the airfoil configuration and is an interchangeable module with variants in cooling and/or provided materials and/or compounding in line with the different operating regimes of the turbomachine.
One of the most important aspects of the invention provides at least one outer shell and, according to operative requirements, at least one intermediate shell for modular variants of the original rotor blade or guide vane airfoil. Function of the blade airfoil core structure is to carry mechanical load from the blade airfoil module. In order to protect the airfoil core structure against high temperature and thermal deformation, an outer and an intermediate shell may be introduced.
Accordingly, the intermediate shell is in any case an optional arrangement. It may be required as compensator for different thermal expansion of outer shell and inner core structure (spar). The outer shell is joined to the optional intermediate shell or to the spar generally by interference fit and the intermediate shell is also joined to the spar by interference fit.
The mentioned spar as the inner core structure includes a tip cap and includes a cooling configuration for cooling the spar itself and optionally for providing the intermediate shell with cooling media.
In the event of a damage to the flow-charged outer shell, repair involves the replacement of only the damaged blade or vane component instead of a replacement of the entire blade or vane. The modular design facilitates the use of various materials referring to the shell, including materials that are dissimilar. Thus, suitable materials can be selected within the shell components to optimize component life, cooling air usage, aerodynamic performance and costs.
The outer shell, charged by the hot gas flow, can further include a seal provided between a recess and at least one of the radial endings of the outer shell and the outer peripheral surface of the blade airfoil at the radial end. As a result, hot gas infiltration or cooling air leakage can be excluded, if the outer shell segments are brazed or welded along their radial interface at or nearby the outer peripheral surface so as to close the gaps. Alternatively, the gaps can be filled with a compliant insert or other seal (rope seal, tongue and groove seal, sliding dove-tail, etc.) to prevent hot gas ingress and migration through the gaps. In all cases, the interchangeability of the single shell or shell components can be maintained.
The gap or groove of the radial interface of the single shell components can be filled with a ceramic rope and/or a cement mixture. An alternative consists in a shrinking process with respect to the shell or shell components on the spar or airfoil core structure.
Referring to the joining process, the shrinking process between the outer shell and inner core structure can be stiffened by local deep welding steps in circumferential direction or by brazing etc. In order to provide a certain elasticity to the joining elements during the shrinking process, the outer shell is divided into two or more parts in the radial direction along the leading edge and trailing edge; thus, pressure side and suction side walls forming separate elements. Along the radial separation plane of the outer shell locking elements are provided.
Additionally, the thickness of the outer shell is inwardly tailored in order to achieve optimal flexibility during the shrinking process. Furthermore, the outer shell can be tailored by different material properties or thickness according to individual conditions of use.
In connection with the inwardly tailored surface of the outer shell, it can be disposed that the external surface of the spar or airfoil sub-structure includes a plurality of grooves or protrusions that are in alignment to each other. Grooves and protrusions form an integral force-locked connection with respect to the inward surface of the outer shell.
At a discontinuity on the surface of the divided outer shell due to an overlap of the assembled parts in the area of leading edge and trailing edge, surface continuity flaps may be provided between the leading edge fastening means on the pressure side part. Likewise, surface continuity flaps may be provided between the trailing edge fastening means on the pressure side part. Ceramic filler may be applied between the continuity flaps to provide an aerodynamically smooth surface over the outer shell. Alternatively, ceramic or CMC filler alone may fill and smooth the indents caused by the fastening means.
The outer shell can be equipped with defined breaking points to make easier the disassembly process.
The assembling of the individual elements or modules is supported by force closure means (force-fit, form-fit arrangement) with a detachable or permanent fixation. Additionally, one or more mechanical fixing means may be inserted into the connection area, wherein the mechanical fixing means are separate parts and they are cast into the connection area with a force-fit or form-fit arrangement.
Different types of seal between airfoil and platform are possible:
1. A “rope seal” as is described for example in U.S. Pat. No. 7,347,424 B2. In this case, there are leakage losses, however.
2. A “brush seal” Also in this case, leakage losses have to be taken into consideration.
3. A temperature-resistant filing material for ensuring a 100%-sealing without leakage losses with simultaneous avoidance of force transmission, for example by means of superplastic material.
According to individual operating conditions of the turbomachine outer shell and core structure may comprise additional means and/or inserts, which are able to resist thermal and chemical stresses, wherein the mentioned means and inserts are in whole or in part interchangeable.
Additionally, a thermally insulating material or a thermal barrier coating (TBC) generally can be applied to various portions of the rotor blade or guide vane assembly.
The main advantages and features of the present invention are as follows:
Thermo-mechanically decoupling of individual parts, especially of outer and, if existing, intermediate shell, and core structure improve the lifetime of the blade or vane compared to an integral design.
Modules with different variants in cooling and/or material configuration can be selected to best fit to the different operating regimes of the gas turbine.
It is possible to introduce an inner core structure (spar) with an extension from the root portion of the blade or vane to its tip.
Furthermore, it is possible to introduce an inner spar comprising an extension from the root portion of the blade or vane to its tip, whereby in the region of the shank the spar has a special contour in concordance with the contour of the opposite form fit elements.
The core structure of the blade or vane shank consists, in radial direction of the airfoil, of an elongated and relatively slim formed portion. The elongated portion extends over the entire height of the footboard mounting part, wherein the foot-side end of the elongated portion having, both sides of the axial expanse of the elongated portion, shapes of teeth, and the bottom of the elongated portion of the shank under-structure may be formed as the final part of the fir-tree-shaped cross-sectional profile. The teeth of the elongated portion of the shank sub-structure may align with the recesses of two-piece footboard mounting elements to provide room for the teeth of the elongated portion. Footboard mounting elements having opposite cracks or clutches correspond in axially direction to the extending contour of the elongated portion of the shank under-structure for the reciprocal axial coupling.
The blade or vane airfoil comprises a one-piece or multi-part outer shell, selected in a manner to optimize component life, cooling usage, aerodynamic performance, and to increase the capabilities of resistance against high temperature stresses and thermal deformation.
Especially, the outer shell can additionally be provided with a selected thermally insulating material or a thermal barrier coating, but this measure is not limited to the outer shell.
The spar or core structure has various passage-ways to supply a cooling medium through the blade or vane.
The rotor blade or guide vane according to the invention is based on convective cooling with selected areas of impingement and/or effusion cooling.
As a matter of principle, the interchangeability of all elements and modules of the blade or vane is assured; repairing and/or reconditioning of all parts with equivalent or, if available, better replacement parts are ensured.
The connection of various elements/modules to one another can be made by friction-locked means, optionally in combination with adherence or the use of a metallic and/or ceramic surface coating, or by a force closure with bolts or rivets, or by HT brazing, active brazing, soldering, shrinking.
In particular, the assembling between outer shell and core structure is preferably effected by means of a shrinking joint.
The platform of the rotor blade or guide vane may be composed of at least two individual parts, which are on the one hand actively connected to the airfoil and on the other hand to the outer shell or shell elements.
A modular design of the rotor blade or guide vane spar or airfoil sub-structure facilitates the use of various shell's materials, including materials that are dissimilar, in accordance with the different operating regimes of the gas turbine.
The modular assembling of the rotor blade or guide vane consists of replaceable and non-replaceable elements, and besides the modular assembling of the rotor blade or guide vane comprising substitutable and non-substitutable elements.
The foregoing and other features of the present invention will become more apparent from the following description and accompanying figures.
The invention shall subsequently be explained in more detail exemplary embodiments in conjunction with the drawing. In the drawing:
a, b shows an assembling solution between an outer shell (
a, b shows an assembling solution between an outer shell (
a, b shows an assembling solution between an outer shell (
In
The inner platform abuts the platforms of neighbouring blades to define a gas passage inner wall for the turbine. An outer not shown thermal shield at the tip of the blade airfoil 118 cooperates again with its neighbours in the manner shown to define the outer wall of the turbine's gas passage.
Cooling passages, which are not shown, extend inside the blade airfoil 110 for cooling purposes, supplied with a cooling medium, particularly cooling air, via a feed hole 117 which is arranged on the shank 114 at the side (see
The shank 114 may be formed with the inner platform 115 or the platform 115 may be formed separately and joined thereto and projects in a circumferential direction to abut against the inner platform in the adjacent rotor blade in the rotor disk (not shown). A seal (not shown) may be mounted between platforms of adjacent rotor blades to minimize or eliminate leakage around the individual rotor blades.
The tip 118 of the rotor blade 100 may be sealed by cap 240 that may be formed integrally with the spar 210, or may be a separate piece that is joined to the top end of the spar 210. The outer shell 220 extends over the surface of the spar 210 between the platform 115 and the tip 240.
The outer shell 220 defines a pressure and suction side, a leading 112 and trailing edge 113 (see
As shown in
According to
Furthermore, the intermediate shell 230 provides additional protection to the spar 210 in case of damage of the outer shell 220. Basically, the intermediate shell 230 is an interchangeable module with variants in cooling and/or material configurations adapted to the different operating regimes of the gas turbine. If several superimposed shells are provided, they may be built with or without spaces between each other.
Outer shell according to
Outer shell 220 is interchangeable, consumable, pre-fabricated, and may be made of a single part or multi-piece, and comprising variants in cooling and/or material configurations adapted to the different operating regimes of the gas turbine. Outer shell 220 is joined to an intermediate shell (see
Number | Date | Country | Kind |
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13199004.6 | Dec 2013 | EP | regional |
Filing Document | Filing Date | Country | Kind |
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PCT/EP2014/077627 | 12/12/2014 | WO | 00 |