The present invention relates to advanced concepts for modular Industrial Gas Turbine (IGT) components, which basing on the principle of using best-suited materials for individual sections in every area of IGT components according to the state of the art. It refers to methods for manufacturing of isolation structures resp. panels, cooling structures resp. panels for applying on a component to adapt the final component to a specific function.
Furthermore the present invention relates to i) combined panel structures; ii) special treatment of the various panels which form the combined panel structure; iii) manufacturing concept referring to a structured cooling panel; iv) various IGT components which are assembled by the combined panel structures.
Moreover, the present invention relates to a modular structure assembly and a CMC airfoil fixation on metallic platform with metallic core of the above-mentioned structured IGT components, especially referring to a guide vane or rotor blade airfoil.
The selection of most adequate material being actively connected to environmental, thermal, mechanical and thermo-mechanical load conditions under service exposure. According to this design concept, monolithic ceramic and ceramic matrix composite (CMC) materials are especially beneficial to be applied in high temperature loaded areas, whereas metal alloys are preferentially used mainly in mechanically or moderate thermo-mechanically loaded sections. This principle implicates the utilization of monolithic ceramics and especially CMCs for the generation of GT components or component modules/specific areas in the Turbine & Combustor GT section such as platforms and airfoils, inserts or more generally as liner material.
The approach is also an extension of the new reconditioning/repair concept referring to WO 2014/146829 A1, adapted to higher T applications. As a matter of fact, monolithic ceramic materials and ceramic matrix composites are less prone to thermal degradation effects when exposed to high and very high temperatures (1000-1700°) and cyclic operation regimes. Compared to the use of metallic alloys, which must be protected by an environmental metallic coating combined with a thermal barrier coating (i.e., TBC system), a considerably longer overall component lifetime is enabled.
Ceramic systems (incl. CMCs) might also need environmental and thermal barrier protection coatings made of ceramics, especially in the higher temperature range, but thus allowing operation at temperature levels that metallic alloys cannot sustain or considerably increasing the component lifetime.
Based on this concept, there are mainly two factors, driving the development of monolithic and ceramic composite made bodies for the Turbine Blading and Combustor sections of land based IGT and, however, this is also valid, at least in part, for the aero GTs:
Typical drawbacks of today's standard monolithic ceramic and CMC systems in modular IGT component designs:
Commercially available, as well as in literature described CMC material still suffers from:
Mechanical strength values of CMCs, which are generally at the limit of the design requirements, when considering turbine parts and especially in case of rotating blading (creep loading). Considering a simple functional split (i.e., mechanical and thermal decoupling) of the different component sections, such as, splitting the component into different subcomponents, where certain areas have mainly to sustain the mechanical load and other component sections will have to resist to a high thermal loading (as mentioned in several patents and open literature) is not sufficient. Some specific areas of the shell are in charge of very high temperatures and to non-negligible mechanical loading from the very high gas mass flow, gas pressure and centrifugal load.
Strong anisotropic mechanical and physical CMC material properties. This phenomenon is based on the intrinsic 2D/3D woven microstructure of inorganic fibres within the CMC composite. Especially in the case of a need for thicker material strengths, multiple layer arrangements are unavoidable, additionally increasing the intrinsic inhomogeneity and leading to the risk of local defects or complete delamination in between the individual stacked layers. Additionally to this aspect, the overall risk of increasing porosity raises with the number of layers, which are used to form the final shell or liner section.
Limited creep behaviour, which is mainly driven by the fibre properties contained as reinforcement elements within the CMC microstructure. This peculiarity of ceramic composite material limits even further the design flexibility and subsequent application in areas of combined mechanical and high temperature loading over long operation times.
High thermal gradients (temperature inhomogeneity) around the airfoil are to be expected. Out to the fact that the resulting maximum thermal and mechanical loading can be very localized, this bears a high risk of local damage formation, which might ultimately lead to a complete failure of the CMC system
It is an object of the present invention to summarize the limits and shortcomings of today's monolithic ceramic and CMC sections for IGT components, as well as the correlated modular component design scenarios. For this purpose, the following critical aspects have to be overcome:
The inventive object referring to an embodiment according to the independent claim.
Furthermore, the object of the present invention concerns a method for assembling a guide vane or a rotor blade of a turbomachine on the basic of a modular structure, wherein at least the airfoil-shells, liners and other components comprising at least one isolation panel prepared by a manufacturing process according to one or more of the attached claims of the present description.
Using the example of a guide vane assembly of a turbomachine on the basis of a modular structure, this guide vane comprises at least one airfoil, an inner platform, an outer platform, wherein the guide vane airfoil and/or platforms have at its one ending provisions for the purpose of a connection of the guide vane elements among each other.
The connections of guide vane elements among each other, especially manufacturing according to one or more of the attached claims of the present description, are configured as a detachable, permanent or semi-permanent fixation with respect to the radial or quasi-radial extension of the airfoil compared to the rotor axis of the turbomachine.
The assembling of the airfoil with respect to at least one platform is based on a forcefit and/or a form-fit connection, or the assembling of the airfoil with respect to at least one platform is based on the use of a metallic and/or ceramic fitting surface, or the assembling of the airfoil with respect to at least one platform is based on force closure means with a detachable, permanent or semi-permanent fixation.
At least the guide vane airfoil or an alternative base structure of the airfoil comprise at least one flow-charged outer hot gas path liner, which encases at least one part of the guide vane airfoil, wherein the flow-charged outer hot gas path liner is connected to the guide vane airfoil or alternative base structure of the airfoil by using a shrinking joint.
Moreover, at least the guide vane airfoil or an alternative base structure of the airfoil comprises at least one flow-charged outer hot gas path liner, which encases at least a part of the guide vane airfoil. The flow-charged outer hot gas path liner is connected with respect to the guide vane airfoil or alternative base structure of the airfoil by using a shrinking joint.
Furthermore, at least the guide vane airfoil or an alternative base structure of the airfoil comprises at least one flow-charged outer hot gas path liner, which encases at least one part of the defined guide vane airfoil, wherein the platforms comprise at least one insert element or mechanical interlock and/or additional thermal barrier coating along thermal stress areas.
Additionally, at least the guide vane airfoil or an alternative base structure of the airfoil comprises at least one flow-charged outer hot gas path liner, which encases at least one part of the defined guide vane airfoil, wherein platforms and/or airfoil and/or airfoil carrier and/or outer hot gas path liner comprise at least one insert element and/or mechanical interlock along or within the thermal stress areas.
The same or similar considerations apply also if it concerns a rotor blade or a liner and associated structures and components.
The method for assembling a guide vane or rotor blade is characterized in that the insert element(s) on the basic of an isolation panel according to one or more of the attached claims of the present description and/or mechanical interlock, forming the respective flow-charged zone, are inserted at least in a force-fitting manner into appropriately designed recesses or in the manner of a push loading drawer with additional fixing means.
As an example of assembling a rotor blade, following procedure is possible: Rotor blade elements comprising at least one rotor blade airfoil, at least one footboard mounting part, whereas rotor blade elements having at its one endings means for the purpose of an interchangeable connection among each other. At least the airfoil is manufactured according one or more of the attached claims of the present description. The connection of the airfoil with respect to other elements based on a fixation in radially or quasi-radially extension compared to the axis of the gas turbine, whereas the assembling of the blade airfoil in connection with the footboard mounting part based on a friction-locked bonding actuated by adherence interconnecting. Alternatively, the assembling of the blade airfoil in connection with the footboard mounting part based on the use of a metallic and/or ceramic surface fixing rotor blade elements to each other. Or the assembling of the blade airfoil in connection with the footboard mounting part based on closure means with a detachable, permanent or semi-permanent fixation, whereas the footboard mounting part consisting of at least twofolded elements. The assembly of separated footboard mounting parts with respect to a foot-side elongated portion of the rotor blade airfoil is conducted with a reciprocal axially guided coupling, whereas the footboard mounting parts (120, 130) having axially opposite cracks or clutches corresponding to the axially extending contour of the elongated portion of the shank under-structure. The axially extending contour of the elongated portion of the shank under structure corresponding approximately to the axially inflow plane of the airfoil.
Additionally, the main target is to decouple thermal and mechanical load referring to a concept of modules and joints in connection with a CMC airfoil, especially referring to a guide vane or rotor blade airfoil.
Fundamentals items for CMC airfoil shell integration to the part based on:
Definitions: a panel consists of a single or multi-plies (=multi-layered) structure of tissue fibres with a defined arrangement. Each layer of a multi-plies structure can be made of a different fibre arrangement, orientation, architecture or geometric consistency, for example size.
A distinction is imposing:
Depending on the needs, a coating can be added by various different processes (thermal spraying, dipping, CVD, etc.) on top of the internal or external surface of the panel [system 3, panel 3, see for example
Depending on the needs, the arrangement can be (see
4. Other combinations can be realised depending on the specific application. The layer-up order of the panels can also be different. The combined structure panels used can be combinations of the different panel systems presented above, namely:
The panels, as single panel or as combined structure panels, can be of complex 3D geometries such a gas turbine rotor blade or vane airfoil.
Moreover, the term (expression) “ceramic textile” is a synonym to “tissue”; the more generic “ceramic fabric” could be used here. The subsequent use of the various terms should be viewed in this context.
The manufacturing steps are using the standard CMC multi-plies production process, in which are integrated some steps or special tissues or items in order to obtain the targeted features.
The advantage of the method is the generation of complex panels (for adaptation of the final product to a specific function) with multiple systems configuration using one single drying and one single sintering step and minimizing (or completely eliminating) the rework steps such as post-machining, surface rework or coating application. The following methods are preferably used as a basis:
Manufacturing concept referring to isolation structures/panels comprising the following operations a)-h):
a) Cutting of a desized 2D ceramic tissue in the right size and shape for the application. Desizing means a process to remove the fibres coating used for manufacturing of the ceramic tissues. b) Slurry infiltration in the tissue by, preferably by knife blade coating (this level open also other slurry infiltration methods) c) Laminating on mould of a single layer or of a multi-layer according to the panel: This method operation comprising at least three steps, namely i) an application of one to n-layers of standand ceramic tissue (with arranged fibres, e.g., un-directionally arranged (UD), or woven fibres); and ii) different layers in the multi-layered panel do not have necessarily to have the same fibre orientation or weaving architecture, and iii) slurry infiltration in the tissue by, e.g., knife blade coating, after each single layer. The three mentioned steps do not necessarily take place in common. d) Laminating on top of multi-layer according to the system 1 (i.e., panel 1) of a single layer or of a multi-layer, according to the system 3 (i.e. panel 3): This method operation comprising at least two steps, namely i) an application of one to n-layers of ceramic felt, and ii) a slurry infiltration in the tissue (only partial: only outer surfaces of the tissue in order to bind it to the CMC multiplies panels, or fully impregnated) by, e.g., knife blade coating, after each single layer. The two mentioned steps do not necessarily take place in common. e) Optional: Pins application (concept with inserted pins) in order to generate straight cooling paths through a part of the thickness or through the full thickness of the multiple panel structure (i.e., cooling air holes, CAH): This method operation comprising at least three steps, namely pins can be i) permanent metallic pins with a ceramic layer coating to avoid attachment of matrix to the pin and too strong oxidation of the pin during sintering; and ii) permanent ceramic pins that can be easily removed after sintering; and iii) pin that will be eliminated during the sintering process via the heat treatment (e.g., carbon pins) leaving the holes structure intact. Additionally, pins are applied by sliding them through a part of thickness or through the whole thickness of the multiple panel structure. In the latter case, in order to facilitate the positioning of the pins, the mould underneath can be provided with positioning hole in which the pins are fit into with the appropriated position and angle. The pins are inserted in-between the tissue fibre bundles in order to avoid any damages of the ceramic fibres during the processing and later removal of the pins. All mentioned steps do not necessarily take place in common. e) Drying. f) De-moulding. g) Sintering the whole structure in one-step in order to finalise the component specific areas or component module. h) Only in case of optional operation e) is made: Remove pins avoiding to damage the fibre bundle surrounding them. Removal technique will depend on the type of pins used [see various steps under e)]. i) Finishing, namely using of i) post-machine, and/or ii) surface smoothening/rework, and/or iii) coating application, and/or other procedures, if necessary.
Elucidating: The knife blade coating is the one used method. This is effectively making more sense to keep infiltration method as general step and mention knife blade coating as an example. Other methods are, e.g., pressure infiltration, pre-processing, electrophoretic deposition, etc.
The initial system applied on the mould must not be mandatorily the system 1. Nevertheless, the initial system can also be system 2 or 3 as described in various figures, wherein the various combinations are described and will depend on the final application or feature targeted.
Manufacturing concept referring to cooling structures/panels (
Manufacturing concept using tissue with combined fibres architectures (
a) Cutting of a desized 2D ceramic tissue in the right size and shape for the application. Desizing in this context means a process to remove the fibres coating used for manufacturing of the ceramic tissues. b) Slurry infiltration in the tissue by, e.g., knife blade coating. c) Laminating on mould of a single layer or of a multi-layer according to system 2 (i.e., panel 2): This method operation comprising at least three steps, namely i) an application of one to n-layers of combined fibre architecture ceramic tissue (see
The initial system applied on the mould must not be mandatorily the system 2. But can also be system 1 or 3 as described in various figures, where the various combinations are described and will depend on the final application or targeted feature.
Manufacturing concept using tissues with combined fibres architectures comprising the following operations a)-j): (
a) Cutting of a desized 2D ceramic tissue in the right size and shape for the application. Desizing in this context means a process to remove the fibres coating used for manufacturing of the ceramic tissues. b) Slurry infiltration in the tissue by, e.g., knife blade coating. c) Laminating on mould of a single layer or of a multi-layer according to system 2 (i.e., panel 2), comprising at least the followings steps: i) an application of one to n-layers of woven tissues with integrated cooling holes structure (see
h) Sintering the whole structure in one-step in order to finalise the component specific areas or component module, namely: During the sintering process, the “sacrificial” fibres (i.e., black fibres depicted in
The initial system applied on the mould must not be mandatorily the system 2. Nevertheless, the initial system can also be system 1 or 3 as described in various figures, where the various combinations are described and will depend on the final application or feature targeted.
Referring to intermediate lawyer the following aspects should be highlighted: In this case, two main functions of the intermediate layer are significant:
In this context, reference to the prior art is made, namely referring to DE 10 201 3110381 A1 and U.S. Pat. No. 8,267,659 B2 where the features (that could correspond to an intermediate layer) are used as spacer between the support structure (e.g., metallic airfoil core) and the CMC shell.
Manufacturing concept referring to an intermediate layer comprising the following operations a)-k):
a) Cutting of a desized 2D ceramic tissue in the right size and shape for the application. Desizing in this context means a process to remove the fibres coating used for manufacturing of the ceramic tissues. b) Slurry infiltration in the tissue by, e.g., knife blade coating. c) Laminating on mould of a single layer or of a multi-layer system 1, 2 or 3. d) Laminating on top of system 1, 2 or 3a single layer or of a multi-layer system 1, 2 or 3. e) Repeating operations c) and d) until the targeted CMC structure is reached. f) Drying. g) De-moulding. h) Combining an “undulated” structure made of CMC (see
i) Drying if operation h) lit. c. is carried out. j) Sintering the whole structure in one step in order to finalise the component specific areas or component module. k) Finishing, namely using of i) post-machine, and/or ii) surface smoothening/rework, and/or iii) coating application, and/or other procedures.
Cooling air holes generation using pins can be added to the above manufacturing sequence if needed.
Flexible Manufacturing concept with reinforcement inserts and insert integration:
The insert can be an airfoil Leading Edge (LE) area, or a Trailing Edge (TE) area, or other area, including platforms that are strongly solicited from a thermal and mechanical loading point of view.
Pre-impregnated CMC panels (AF/SS & AF/PS=see
Together with an impregnation, a slurry infiltration by various methods (for example combing method) can also be applied.
The final part with insert at the LE and bound area at C+D points is shown in
The following operations are made a)-m):
a) Cutting of a desized 2D ceramic tissue in the right size and shape for the application. Desizing in this context means a process to remove the fibres coating used for manufacturing of the ceramic tissues. b) Slurry infiltration in the tissue by, e.g., knife blade coating in the areas A to C and B to D. c) Laminating on mould of a multi-layer system 1, 2 or 3, wherein LE (Leading edge) area, within the meaning of the above specification, remains thinner taking into account the thickness of the insert. d) Optional: Laminating on top of system 1, 2 or 3a single layer or of a multi-layer system 1, 2 or 3 in the areas A to C and B to D. e) Repeating operations c) and d) until the targeted CMC structure is reached. f) Drying. g) De-moulding. h) Enveloping the insert, which has been pre-manufactured and therefore it is readily available to be integrated in the envelope for the final manufacturing steps with the pre-prepared panel system as shown in
i) Slurry infiltration in the tissue by, e.g., knife blade coating in the LE area (points A to B in
Cooling air holes generation using pins can be added to the above manufacturing sequence if needed.
The present invention is going now to be explained more closely by means of different embodiments and with reference to the attached drawings.
Figures of this description show various embodiment of single and/or multi-plies CMC panels provided for system arrangements. Fundamentally, the CMC panel can be designed with individualized fibre structure in accordance with the operational requirements. A certain percentage of the fibres may have differentiated diameters, which are intended to mainly carry the mechanical load (in the case of the larger diameters) and thermal stresses during the flow-applied operation.
Other combinations can be realised depending on the specific application. The layer-up order of the panels can also be different. The panels can be made of complex 3D geometries such as a gas turbine rotor blade/vane airfoil (see the examples under
The mentioned CMC zones 20, 40 can consist of a laminate structure, such that an appropriate bond between the single intermediate layers (different tissue plies) is achieved. Furthermore, the zones can be formed by a multiple sandwich structure. “Laminate structure” means the technique of manufacturing a material in multiple layers, so that the composite material achieves improved strength, stability, sound insulation, appearance or other properties from the use of differing materials. A laminate structure is usually permanently assembled by heat, pressure, welding, or adhesives
Moreover, the mentioned ceramic felt 30 between the CMC zones 20, 40, likewise build-up of 2D/3D tissue structure with thinner fibres, serves to fix the ceramic matrix to the overall fibre substructure. The fibres of the ceramic felt can be differently woven using the same or different materials, within the ceramic felt and on each side of the panels comprises both first and second fibre materials. Any stacking-sequence of different woven fibres within the thickness of the panel arrangement is also possible.
Additionally, the cooling holes being actively connected to the structured cooling or chaotic running channels within the panel body are designated for convective and/or impingement and/or effusion cooling effects. Furthermore, in some specific configurations or using a specific type of pins, the introduced pins (see
One additional point to mention is the fact that the pins can be inserted through the full thickness of a single panel or of a panel combination. However, it can also be the case that the pins are inserted only partially through the thickness of the panel combination, not passing through the whole thickness.
Complementary fibres 60 (black) consist of carbon fibres. The “black” fibres are “sacrificial fibres that will burn out during the sintering process leaving a negative architecture forming the cooling structure. Both fibres can be differently woven using the same or different materials. The resulting architecture having a rectangular or quasirectangular weaving, or an oblique or quasi-oblique, or non-rectangular angulation weaving. Furthermore, the architecture can be designated as a sinusoidal or quasisinusoidal interdigitated weaving. Any stacking-sequence of different woven fibres within the thickness of the panel arrangement is also possible.
A practical result of a manufacturing according to
In this context using a special fabric (see definition under “summery of the invention”) where the cooling structure (CAHs) is directly woven in within a ceramic fabric and applying as a layer. The performing are as follows:
Referring to
The additional use of a heat and oxidation resistant flexible layer (
Two ceramic tissues interwoven at one extremity in order to enable a connection at, e.g., the TE, which is not only relying on gluing/brazing methods.
Accordingly, each guide vane provides a radial outer platform 200, an airfoil 100 and a radial inner platform 300. The radial outer platform contains mounting hooks 201, 202 that are inserted into mounting grooves of the stator component of the first turbine stage (not shown). The inner platform 300 of the guide vane, typically, encloses a gap with the rotor liner through which a purge flow of cooling medium can be injected into the hot gas flow within the gas turbine. In the same way, a purge flow of cooling medium is injected through a gap, which is enclosed by parts of the stator component, the upstream edge of the outer platform 200 of the guide vane and the outer combustor liner, also called stator liner. Generally, downstream of the outer platform 200a heat shield (not shown) is mounted inside of the stator component which prevents overheating of the inner faced areas of the stator component in the same way as in case of the outer platform 200.
Generally, the means for the purpose of an interchangeable connection of the guide vane elements, namely between airfoil, inner platform, outer platform and optionally flow carrier comprising reciprocal lugs or recesses based on a friction-locked bonding or permanent connection or fixing.
The flow-applied shell module encases integrally or partially the outer contour of the based guide vane airfoil of the guide vane according to aerodynamic requirements. The partial shell structure is actively connected to the leading edge of the based airfoil of the guide vane, wherein the outer contour of the based airfoil consists of an independent flow-charged part, being actively connected to the leading edge of the airfoil of the guide vane. The flow-charged shell structure encases integrally the outer contour of the based guide vane airfoil, complying with aerodynamic final aims of the vane, or the flow-charged shell structure encases partially the outer contour of the based airfoil in the flow direction of the working medium of the gas turbine, complying with aerodynamic final aims of the guide vane. According to an additional embodiment the based guide vane airfoil comprises inside a supplementary body formed by the configuration of a spar. In place of the based guide vane airfoil can be made a spar as substructure. The shell structure may be formed by the form of an integrally or segmented body. The first shell structure comprises internally a second or intermediate non-flow-charged or partially flow-charged shell structure, complying with aerodynamic final aims of the vane. The two shell liners are adjacent or have an intermediate distance from one another. When the first flow-charged shell structure encases integrally the outer contour of the guide vane airfoil, this shell structure comprises at least two bodies forming completely or partially the outer contour of the based guide vane airfoil. The mentioned bodies, forming completely or partially the outer shell structure, are brazed or welded along their radial interface, and they have radial or quasi-radial gaps, which are filled with a seal and/or ceramic material. The outer shell is inter-changeable, consumable, pre-fabricated, single or multi-piece with radial or circumferential patches or uses with respect to the sub-structure of the guide vane airfoil a shrinking joint.
Furthermore, the intermediate shell or shells are parts of an optional assembly. The mentioned shell(s) are inter-changeable, pre-fabricated, arranged as single or multi multi-piece with radial or circumferential patches, uncooled or cooled (convective, film, effusion, impingement cooling), fabricated as compensator for different thermal expansion of outer shell and spar, and with a cooling shirt with respect to different cooling configurations for optimization operational requirements. The spar as sub-structure of the guide vane airfoil or of the shell assembly is interchangeable, pre-fab heated or various manufactured, single or multi-piece, uncooled or cooled using convective, film, effusion, impingement cooling, having a web structure for cooling or stiffness improvement.
Normally, the platforms 200, 300 and the guide vane airfoil are no consumable parts. In contrast, the mentioned sealing and liners are consumable parts. The airfoil carrier may be consumable, depending on costs. The airfoil carrier 220 is cast, machined or forged comprising additionally additive features with internal local web structure for cooling or stiffness improvements. Furthermore, the airfoil carrier comprises flexible cooling configurations for adjustment to operational requirements, like base-load, peak-mode, partial load of the gas turbine.
The outer platform 200 is cast, forged or manufactured in metal sheet or plate. The outer platform is consumable in relation to predetermined cycles, and frequently replaced at specified maintenance periods, and may be mechanically decoupled from the guide vane airfoil, wherein the outer platform may be supplementary mechanically connected to the airfoil carrier, using force closure elements, namely bolts. The outer platform may be coated with CMC or ceramic materials or may be manufactured by an isolation panel according to the attached claims.
The inner platform 300 is cast, forged or manufactured in metal sheet or plate. The outer platform is consumable, is replaced at specified maintenance periods, and may be mechanically decoupled from the guide vane airfoil, wherein the inner platform may be supplementary mechanically connected to the airfoil carrier, using force closure elements, namely bolts. The inner platform may be coated with CMC or ceramic materials or may be manufactured by an isolation panel according to the attached claims.
Number | Date | Country | Kind |
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15176309.1 | Jul 2015 | EP | regional |