Integrated nozzle and diaphragm with optimized internal vane thickness

Abstract

The present invention provides a blade for a turbine system. The blade includes: an internal chamber configured to receive a flow of cooling fluid; a variable thickness wall adjacent to the internal chamber; and an impingement plate separating the variable thickness wall from the internal chamber, the impingement plate including a plurality of openings for directing the cooling fluid into an impingement chamber and against the variable thickness wall, wherein the impingement plate is configured to follow the contour of the variable thickness wall.

Description

BACKGROUND OF THE DISCLOSURE

[0001] The present disclosure relates generally to gas turbine systems and, more particularly, to an integrated nozzle and diaphragm having optimized internal vane thicknesses.

[0002] Gas turbine systems are an example of turbines that are widely used in fields such as power generation. Conventional gas turbine systems typically include a compressor section, a combustor section, and a turbine section. During operation of a gas turbine system, various components in the gas turbine system, such as nozzle vanes, turbine blades, and shroud segments, are subjected to high-temperature gas flows and associated thermo-mechanical forces, which can lead to component failure. SUMMARY OF THE DISCLOSURE

[0003] A first embodiment relates to a vane of a turbine system. The vane includes: an internal chamber configured to receive a flow of cooling fluid; a variable-thickness wall adjacent to the internal chamber; and an impingement plate separating the variable-thickness wall from the internal chamber, the impingement plate including a plurality of openings for directing the cooling fluid into an impingement chamber and against the variable-thickness wall, wherein the impingement plate is configured to follow the contour of the variable-thickness wall.

[0004] Another embodiment provides a nozzle segment for a gas turbine system. The nozzle segment includes an integrated nozzle and diaphragm, the nozzle including at least one vane, each vane including: an internal chamber configured to receive a flow of cooling fluid; a variable-thickness wall adjacent to the internal chamber; and an impingement plate separating the variable-thickness wall from the internal chamber, the impingement plate including a plurality of openings for directing the cooling fluid into an impingement chamber and against the variable-thickness wall, wherein the impingement plate is configured to follow the contour of the variable-thickness wall.

[0005] Another embodiment relates to a method for optimizing a vane of a gas turbine system, the method including: determining an operating force on the vane; and changing a thickness of a wall of the vane based on the operating force on the vane; wherein the thickness of the wall is greater in regions of higher operating force on the vane.

[0006] Exemplary aspects of the present disclosure address the problems described herein and / or other problems not discussed. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] These and other features of the present disclosure will be more readily understood from the following detailed description of various aspects of the present disclosure in conjunction with the drawings depicting embodiments of the present disclosure.

[0008] Figure 1 A schematic diagram of a gas turbine system according to an embodiment is depicted.

[0009] Figure 2Depicts a side view of a portion of a turbine section of a gas turbine system according to an embodiment.

[0010] Figure 3 Depicts an isometric view of a nozzle section including a nozzle with an integrated diaphragm according to an embodiment.

[0011] Figure 4 Depicts according to an embodiment along Figure 3 A cross-sectional view of a nozzle vane taken along line 4-4 in

[0012] Figure 5 Depicts according to an embodiment along Figure 4 A cross-sectional view of a portion of a nozzle vane taken along line 5-5 in

[0013] Figure 6 Depicts according to an embodiment along Figure 5 A cross-sectional view of a portion of a nozzle vane taken along line 6-6 in

[0014] Figure 7 Depicts according to an embodiment along Figure 4 A cross-sectional view of the leading edge portion of a nozzle vane taken along line 7-7 in

[0015] Figure 8 Depicts a cross-sectional view of a nozzle section according to an embodiment.

[0016] Figure 9 Depicts a flowchart of a process for optimizing the wall thickness in a nozzle vane according to an embodiment.

[0017] Figure 10 Depicts a block diagram of an additive manufacturing process according to an embodiment of the present disclosure, the additive manufacturing process including a non-transitory computer-readable storage medium storing code representing an object.

[0018] It should be noted that the drawings of the present disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of the present disclosure and should not therefore be considered as limiting the scope of the present disclosure. In the drawings, like numbers represent like elements between the drawings. Detailed Description

[0019] Reference will now be made in detail to representative embodiments illustrated in the accompanying drawings. It should be understood that the following description is not intended to limit the embodiments to one embodiment. Instead, it is intended to cover alternatives, modifications, and equivalents that may be included within the spirit and scope of the described embodiments as defined by the appended claims.

[0020] First, in order to clearly describe the present disclosure, it will be necessary to select certain terms when referring to and describing related machine components within the scope of the present disclosure. In doing so, where possible, common industry terms will be used and adopted in a manner consistent with their accepted meanings. Unless otherwise specified, such terms should be given a broad interpretation consistent with the context of the present application and the scope of the appended claims. Those of ordinary skill in the art will appreciate that several different or overlapping terms may typically be used to refer to a particular component. An object that may be described herein as a single part may include multiple components and in another context be referred to as being composed of multiple components. Alternatively, an object that may be described herein as including multiple components may elsewhere be referred to as a single part.

[0021] In addition, several descriptive terms may be regularly used herein, and it should prove helpful to define these terms at the beginning of this section. Unless otherwise specified, these terms and their definitions are as follows. As used herein, "downstream" and "upstream" are terms indicating a direction relative to the fluid flow direction, such as the working fluid passing through a turbine, or for example the air flow through a burner or the coolant through a component system of a turbine. The term "downstream" corresponds to the fluid flow direction, and the term "upstream" refers to the direction opposite to the flow. In the absence of any additional particularity, the terms "front" and "rear" refer to directions, where "front" refers to the front end or compressor end of the engine, and "rear" refers to the rear end or turbine end of the engine. Additionally, the terms "frontward" and "backward" may be used and / or understood separately as being descriptively similar to the terms "front" and "rear". Generally, it is necessary to describe parts at different radial, axial, and / or circumferential positions. The "A" axis represents the axial orientation. As used herein, the terms "axial" and / or "axially" refer to the relative position / direction of an object along the axis A, which is substantially parallel to the rotational axis of the gas turbine system (particularly the rotor section). As further used herein, the terms "radial" and / or "radially" refer to the relative position / direction of an object along the direction "R" (see Figure 1 ), which is substantially perpendicular to the axis A and intersects the axis A at only one position. Finally, the term "circumferential" refers to movement or position around the axis A (e.g., the direction "C").

[0022] In various embodiments, components described as being "fluidly coupled" or "in fluid communication" with each other may be joined along one or more interfaces. In some embodiments, these interfaces may include joints between different components, and in other cases, these interfaces may include rigid and / or integrally formed interconnections. That is, in some cases, components that are "joined" to each other may be formed simultaneously to define a single continuous member. However, in other embodiments, these joined components may be formed as separate members and subsequently joined by known processes (e.g., fastening, ultrasonic welding, bonding).

[0023] Where an element or layer is referred to as being "on another element", "engaged to another element", "connected to another element", or "coupled to another element", it may be directly on, engaged to, connected to, or coupled to the other element, or intervening elements may be present. In contrast, where an element is referred to as being "directly on another element", "directly engaged to another element", "directly connected to another element", or "directly coupled to another element", no intervening element or layer may be present. Other words used to describe the relationship between elements should be interpreted in a similar manner (e.g., "between" versus "directly between", "adjacent" versus "directly adjacent", etc.). As used herein, the term "and / or" includes any and all combinations of one or more of the associated listed items.

[0024] Figure 1 A schematic diagram of a gas turbine system 10 according to various embodiments is shown. As shown, the gas turbine system 10 includes a compressor section 12 that compresses an incoming air flow 14 and delivers a compressed air flow 16 to a combustor section 18. The combustor section 18 mixes the compressed air flow 16 with a pressurized fuel supply 20 and ignites the mixture to form a combustion gas flow 22. Although only a single combustor section 18 is shown, the gas turbine system 10 may include any number of combustor sections 18. The combustion gas flow 22 is then delivered to a turbine section 24. The combustion gas flow 22 drives the turbine section 24 to produce mechanical work. The mechanical work generated in the turbine section 24 may drive the compressor section 12 via a shaft 26 and may be used to drive an external load 28 (such as a generator, etc.).

[0025] Figure 2 is a cross-sectional side view of a portion of the turbine section 24 of the gas turbine system 10 that may incorporate various embodiments disclosed herein. As Figure 2 shown, the turbine section 24 may include a plurality of turbine stages. For example, the turbine section 24 may include a first turbine stage 30A, a second turbine stage 30B, and a third turbine stage 30C. However, the turbine section 24 may include more or fewer turbine stages as needed or desired.

[0026] Each turbine stage 30A - 30C may sequentially include corresponding turbine nozzle rows (hereinafter referred to as "nozzles") 32A, 32B, and 32C and corresponding turbine blade rows (hereinafter referred to as "blades") 34A, 34B, and 34C that are axially spaced apart along the shaft 26 ( Figure 1 ). During operation of the gas turbine system 10, each of the nozzles 32A - 32C remains stationary relative to the blades 34A - 34C. Each of the nozzle rows 32B, 32C is respectively coupled to or integrally formed with a corresponding diaphragm 42B, 42C. The turbine shrouds 44A, turbine shrouds 44B, and turbine shrouds 44C circumferentially enclose the corresponding blade rows 34A - 34C. The housing or casing 36 circumferentially surrounds each stage 30A - 30C of the nozzles 32A - 32C and the blades 34A - 34C.

[0027] The nozzles 32A - 32C and the blades 34A - 34C extract kinetic energy and / or thermal energy from the combustion gas 22. This energy extraction drives the shaft 26. The combustion gas 22 then exits the turbine section 24 and the gas turbine system 10. As will be discussed in more detail below, a portion of the compressed air 16 can be used as a cooling fluid for cooling various components of the turbine section 24, particularly including the nozzles 32A - 32C and the blades 34A - 34C.

[0028] Figure 3 is an isometric view of a nozzle section 46 including an integrated nozzle 32 and diaphragm 42 according to an embodiment. The nozzle 32 and the diaphragm 42 can be formed as a single unit using, for example, an additive manufacturing process. As Figure 3 shown, the nozzle 32 can include an inner wall 48 (which also forms the upper wall of the diaphragm 42) and an outer wall 50 that is radially spaced apart from the inner wall 48. The nozzle 32 can include a pair of vanes 52 that extend from the inner wall 48 across to the outer wall 50. This nozzle configuration is commonly referred to as a doublet in the industry. However, the nozzle 32 can have only one vane 52 (i.e., a singlet) or three (i.e., a triplet) or more vanes 52.

[0029] As Figure 3 shown, the inner wall 48 and the outer wall 50 of the nozzle 32 include various surfaces. More specifically, the inner wall 48 includes a radially outer surface 54 and a radially inner surface 56 that is radially inwardly positioned from the radially outer surface 54. Similarly, the outer wall 50 includes a radially inner surface 58 and a radially outer surface 60 that is radially outwardly oriented from the radially inner surface 58. The radially inner surface 58 of the outer wall 50 and the radially outer surface 54 of the inner wall 48 respectively define an inner radial flow boundary and an outer radial flow boundary for the combustion gas 22 to flow through the turbine section 24 ( Figure 1 ).

[0030] As described above, two vanes 52A, 52B (commonly referred to herein as vanes 52) extend from the inner wall 48 to the outer wall 50 of the nozzle 32. As Figure 3 shown, the body of each vane 52 includes a leading edge 62, a trailing edge 64, a pressure sidewall 66, and opposing suction sidewalls 68 extending from the leading edge 62 to the trailing edge 64.

[0031] Cooling fluid (such as pressurized cooling air 70 discharged from the compressor section 12 of the turbine system 10 ( Figure 1 )) can be directed into one or more internal chambers 72 formed within each vane 52. The cooling air 70 can be used to cool (e.g., by impingement cooling, convective cooling, film cooling, etc.) various internal and external portions of the vane 52.

[0032] Various portions of the vane 52 of the nozzle 32 (including the leading edge 62 and portions known in the art as high-c regions) may be subjected to high temperatures and high mechanical forces during operation of the turbine system 10 ( Figure 1 ), which can result in a shortened operating life of the nozzle 32. According to an embodiment, the operating life of the nozzle 32 can be increased, for example, by preferentially varying the thickness of various portions of the vane 52 (e.g., thicker in regions subjected to higher forces and thinner in regions subjected to lower forces) and by providing a contoured impingement plate 76 within the vane 52 ( Figure 4 ).

[0033] Figure 4 Depicts a cross-sectional view of the vanes 52A, 52B of the nozzle 32 taken along line 4-4 in Figure 3 . As shown, each vane 52A, 52B includes at least one internal chamber 72 configured to receive a flow of cooling air 70. The cooling air 70 is shown flowing radially downward (i.e., into the page) into the chamber 72, but other flow directions can be used. The internal chambers 72 of the vanes 52A, 52B can have different configurations as shown (e.g., the wall thicknesses of the vanes 52A, 52B can be different) or can have similar configurations. In the Figure 4 non-limiting example shown, the flow of cooling air 70 is directed radially downward toward the diaphragm 42 ( Figure 3 ) into the internal chamber 72.

[0034] The impingement plate 76 can extend continuously around the internal chamber 72 as shown, or can include a plurality of separate impingement plate segments. Generally, the impingement plate 76 directs the cooling air 70 from the internal chamber 72 of each vane 52 through a plurality of openings 78 formed through the impingement plate 76 ( Figure 5), and enters the impingement chamber 80. After entering the chamber 80, the cooling air 70 impinges on the respective inner walls of the blade 52 (e.g., the pressure side wall 66 and the suction side wall 68 of the blade 52), thereby providing impingement cooling. The cooling air 70 can be directed from the chamber 80 to other internal / external parts of the blade 52 to provide additional cooling to the blade 52.

[0035] The thickness of one or more parts of the blade 52 (e.g., the pressure side wall 66, the suction side wall 68, the leading edge wall 82, the trailing edge wall 84, etc.) can preferably vary according to the expected (or estimated) operating forces. The operating forces on the blade 52 can be determined, for example, via computer modeling, physical testing, or other suitable analytical techniques. This can include, for example, modeling and analyzing the operating forces using engineering simulation tools, modifying one or more thicknesses, and re-running the analysis. In Figure 5 , for example, the pressure side wall 66 of the blade 52 has been formed to have a substantially uniform thickness because the expected operating forces are relatively uniform along the length of the pressure side wall 66. However, the radially outward portion 86 of the suction side wall 68 (e.g., in the high-c region) is formed to have a thickness greater than the thickness of the radially inward portion 88 of the suction side wall 68 because the expected operating forces are greater at the radially outward portion 86 of the suction side wall 68. Generally, the thickness of one or more parts of the blade 52 is proportional to the expected operating forces. For example, according to an embodiment, the thickness of the suction side wall 68 in the high-c region (e.g., see Figure 5 ) can be about 1.1 times to about 1.5 times the nominal wall thickness of the blade 52 (e.g., the thickness of the pressure side wall 66). Generally speaking, depending on the expected operating forces, the wall thicknesses of the various parts of the wall of the blade 52 can be in the range of 1.1 times to 2.5 times the nominal wall thickness of the blade 52.

[0036] According to an embodiment, the impingement plate 76 is configured to follow the contour of the inner wall of the blade 52. For example, as Figure 5 and Figure 6 shown, the impingement plate 76 is configured to follow the contours of the inner side walls 90, 92 of the pressure side wall 66 and the suction side wall 68 of the blade 52. Advantageously, by following the contour of the inner wall of the blade 52, the distance D between the impingement plate 76 and the inner wall of the blade 52 (and the impingement cooling provided via the impingement plate 76) can be more accurately controlled (e.g., as compared to an impingement plate insert). The distance D can be substantially constant throughout the blade 52, or it can be variable (e.g., to selectively adjust the resulting impingement cooling).

[0037] A set of support beams 94 can be provided to connect the impact plate 76 to the inner walls 90, 92 of the pressure sidewall 66 and the suction sidewall 68 and to separate the impact plate from the inner walls. The support beams 94 provide several functions. For example, the support beams 94 connect the impact plate to provide structural support / stiffness for the pressure sidewall 66 and the suction sidewall 68, enabling the pressure sidewall 66 and the suction sidewall 68 of the vane 52 to be made thinner. In addition, the support beams 94 maintain and control the distance D between the impact plate 76 and the inner walls 90, 92 of the pressure sidewall 66 and the suction sidewall 68. When using thinner walls, the cooling air 70 can more effectively cool the hot side of the vane 52. For example, instead of making the vane 52 0.2” thick, it can be made 0.1” thick, and thus the impact plate 76 can be made 0.1” thick. The resulting structure has similar stiffness but better cooling efficiency. Therefore, the amount of cooling air 70 required for the cooling structure is reduced and the life of the structure can be extended. This also provides the ability to fine-tune specific locations to be thinner or thicker than the average, thus addressing stress and oxidation issues.

[0038] According to an embodiment, the impact plate 76 can have a uniform thickness T throughout the vane 52 IP ( Figure 6 ). Alternatively, the thickness T of the impact plate 76 IP can vary within the vane 52. For example, the thickness T of the impact plate 76 IP can be inversely proportional to the thickness T of the adjacent inner wall section W ( Figure 6 ). Thus, the impact plate 76 can be thicker adjacent to thinner inner wall sections (e.g., to provide additional structural support) and thinner adjacent to thicker inner wall sections.

[0039] Figure 7 Depicts a cross-sectional view of the leading edge 62 of the nozzle 52 taken along line 7-7 in Figure 4 . During operation, the radially outward section 96 of the leading edge wall 82 of the vane 52 may experience higher forces than the radially inward section 98 of the leading edge wall 82. To that extent, according to an embodiment, the thickness of the leading edge wall 82 can be preferentially varied such that the higher force region (e.g., the radially outward section 96) is thicker than the lower force region (the radially inward section 98).

[0040] In addition to any impingement cooling provided to the inner wall 112 of the leading edge wall 82 of the vane 52 via the impact plate 76, further cooling may be required at the thickened section 96 of the leading edge wall 82. For example, such cooling can be provided by forming channels 100 that extend from the internal chamber 72 to the chamber 80 through the thickened section 96 of the leading edge wall 82.

[0041] Passage 100 is fluidly connected to chamber 80, and thus further fluidly connected to passage 102 formed inside the wall 48 of nozzle 32. After passing through passage 100 from chamber 72 and absorbing heat from the thickened section 96 of leading edge wall 82, cooling air 70 exits chamber 80 and passes through passage 102 into the internal chamber 104 of diaphragm 42. The cooling air 70 passing through passage 102 into chamber 104 can pressurize chamber 104 (e.g., to prevent hot gas from entering chamber 104) and / or can provide cooling to the inner wall 48 of nozzle 32. Another passage 106 can be provided to fluidly connect chamber 72 to the wheel space 108 of turbine section 24( Figure 1 ). Passage 106 can extend from chamber 72 to wheel space 108, for example, through the inner wall 48 of nozzle 32 and the side wall 110 of diaphragm 42. The cooling air 70 flowing through passage 106 can provide cooling to the inner wall 48 of nozzle 32 and the side wall 110 of diaphragm 42, and / or can be used to pressurize wheel space 108 to prevent hot gas from entering wheel space 108.

[0042] Figure 8 A cross-sectional view of nozzle section 46 taken from the leading edge 62 to the trailing edge 64 of blade 52 according to an embodiment is depicted. To this extent, Figure 8 includes the features previously described at the leading edge 62 of blade 52 Figure 7 .

[0043] As Figure 8 shown, impingement plate 76 follows the contour of the inner wall 114 of trailing edge wall 84 along the trailing edge 64 of blade 52. Cooling air 70 passes through a plurality of openings 78 formed in impingement plate 76 and enters chamber 80, where the cooling air 70 impinges on the inner wall 114 of trailing edge wall 84, thereby providing impingement cooling.

[0044] Chamber 80 can be fluidly connected to at least one of trailing edge cooling circuit 116 and passage 118 formed through the inner wall 48 of nozzle 32. After impinging on the inner wall 114 of trailing edge wall 84, cooling air 70 can exit chamber 80 and pass through passage 118 into the internal chamber 104 of diaphragm 42 and / or can exit chamber 80 to trailing edge cooling circuit 116. The cooling air 70 passing through passage 118 into chamber 104 combines with the cooling air 70 passing through passage 102 into chamber 104 to pressurize chamber 104 (e.g., to prevent hot gas from entering chamber 104) and / or to provide cooling to the inner wall 48 of nozzle 32.

[0045] An additional passage 120 can be provided to fluidly connect chamber 72 to the wheel space 108 of turbine section 24( Figure 1)。The passage 120 can extend from the chamber 72 to the wheel space 108, for example, through the inner sidewall 48 of the nozzle 32 and the sidewall 122 of the diaphragm 42. The cooling air 70 flowing through the passage 120 can provide cooling to the inner sidewall 48 of the nozzle 32 and the sidewall 122 of the diaphragm 42, and / or be used to pressurize the wheel space 108 to prevent hot gas from entering the wheel space 108.

[0046] The support beams 94 formed between the impact plate 76 and the inner sidewalls 112, 114 of the leading edge wall 82 and the trailing edge wall 84 provide structural support / stiffness to the leading edge wall 82 and the trailing edge wall 84, which enables the leading edge wall 82 and the trailing edge wall 84 of the blade 52 to be made thinner. In addition, the support beams 94 maintain and control the separation between the impact plate 76 and the inner sidewalls 112, 114 of the leading edge wall 82 and the trailing edge wall 84.

[0047] Although described above with reference to the turbine nozzle section 46, some / all of the embodiments described herein can also be applied to other components of the gas turbine system 10. For example, some / all of the embodiments described herein can be applied to the turbine blade 34, the shroud 44, or other components of the turbine system 10.

[0048] The various components and features of the nozzle section 46 of the present disclosure can be formed using an additive manufacturing process. Advantageously, additive manufacturing enables the design and production of more customizable features (e.g., blades 52 with optimized wall thicknesses, impact plates 76 following a profile, etc.) and more complex features (e.g., support beams 94) to provide better aerodynamics and high-temperature efficiency. Additionally, by using additive manufacturing, the dimensions (e.g., diameter, length) of the various passages (e.g., passages 100, 102, 106, 118, 120) within the blades 52 and the diaphragm 42 of the nozzle section 46 can be customized / optimized to reduce flow losses and increase cooling efficiency. Additionally, features such as the impact plates 76 following a profile and the support beams 94 can be integrally formed with the walls of the blades 52.

[0049] As described above, for example, by preferentially changing the thickness of the various parts (e.g., walls, impact plates, etc.) of the blade 52, the operating life of the nozzle 32 can be extended. Figure 9 A flowchart of a process for optimizing the wall thickness in the blade 52 is provided.

[0050] At process A1, an analysis is performed to determine the expected operating forces on the various parts of the vane 52 (e.g., pressure sidewall 66, suction sidewall 68, leading edge wall 82, trailing edge wall 84, etc.). The analysis can be performed, for example, on the design of the vane 52, a physical model of the vane 52, or the vane 52 itself. At process A2, the design of the vane 52 is modified by preferentially changing the thickness of one or more walls of the vane 52 based on the expected operating forces. Processes A1 and A2 can be repeated on the updated design of the vane 52 as needed (yes at process A3) to further optimize the wall thickness of the vane 52.

[0051] As used herein, additive manufacturing can include any process of producing an object by successively layering materials rather than removing materials (as is the case in conventional processes). Additive manufacturing can form complex geometries without the use of any kind of tool, die, or fixture, and with little or no waste of material. Instead of machining a part from a solid plastic or metal blank (much of which is cut away and discarded), the only material used in additive manufacturing is the material needed to form the part. Additive manufacturing processes can include, but are not limited to: 3D printing, rapid prototyping (RP), direct digital manufacturing (DDM), binder jetting, selective laser melting (SLM), and direct metal laser melting (DMLM). In the current setup, DMLM or SLM has been found to be advantageous.

[0052] To illustrate an example of an additive manufacturing process, Figure 10 A schematic / block diagram of an exemplary computerized additive manufacturing system 900 for generating an object 902 is shown. In this example, system 900 is arranged for DMLM. It should be understood that the general teachings of the present disclosure apply equally to other forms of additive manufacturing. The object 902 is shown as a nozzle segment 46 ( Figures 3 to 7 ). The AM system 900 generally includes a computerized additive manufacturing (AM) control system 904 and an AM printer 906. As will be described, the AM system 900 executes code 920, which includes a set of computer-executable instructions that define the object 902, to physically generate the object 902 using the AM printer 906. Each AM process can use different raw materials in the form of, for example, fine-grained powder, liquid (e.g., polymer), sheet, etc., the stock solution of which can be held in a chamber 910 of the AM printer 906. In this case, the nozzle segment 46 can be made of a material capable of withstanding the gas turbine system 10 (see Figure 1) and made of metal or metal compound in an environment. As shown in the figure, the applicator 912 can form a thin layer of the raw material 914, which spreads out as a blank canvas on the build plate 915 of the AM printer 906, and each successive slice of the final object will be formed based on this blank canvas. In other cases, the applicator 912 can directly apply or print the next layer onto the previous layer defined by, for example, the code 920, such as in the case of using the metal binder jetting process. In the example shown, the laser or electron beam 916 melts the particles for each slice as defined by the code 920, but this may not be necessary in the case of using rapidly solidifying liquid plastics / polymers. Various parts of the AM printer 906 can be moved to accommodate the addition of each new layer. For example, after each layer, the build platform 918 can be lowered, and / or the chamber 910 and / or the applicator 912 can be raised.

[0053] The AM control system 904 is shown to be implemented as computer program code on a computer 930. To this extent, the computer 930 is shown to include a memory 932, a processor 934, an input / output (I / O) interface 936, and a bus 938. In addition, the computer 930 is shown to communicate with external I / O devices / resources 940 and a storage system 942. Generally, the processor 934 executes the computer program code stored in the memory 932 and / or the storage system 942, such as the AM control system 904, under the instructions from the code 920 representing the object 902 described herein. When executing the computer program code, the processor 934 can read and / or write data to / from the memory 932, the storage system 942, the I / O device 940, and / or the AM printer 906. The bus 938 provides a communication link between each component in the computer 930, and the I / O device 940 can include any device that enables a user to interact with the computer 940 (e.g., keyboard, pointing device, display, etc.). The computer 930 only represents various possible combinations of hardware and software. For example, the processor 934 can include a single processing unit or be distributed across one or more locations (e.g., on a client and a server). Similarly, the memory 932 and / or the storage system 942 can reside at one or more physical locations. The memory 932 and / or the storage system 942 can include any combination of various types of non-transitory computer-readable storage media, including magnetic media, optical media, random access memory (RAM), read-only memory (ROM), etc. The computer 930 can include any type of computing device, such as a web server, a desktop computer, a laptop computer, a handheld device, a mobile phone, a pager, a personal digital assistant, etc.

[0054] The additive manufacturing process begins with a non-transitory computer-readable storage medium (e.g., memory 932, storage system 942, etc.) that stores code 920 representative of object 902. For example, code 920 may include a precisely defined 3D model of object 902 and may be generated by any of a variety of well-known computer-aided design (CAD) software systems (such as DesignCAD 3D Max, etc.). In this regard, code 920 may be in any currently known or future-developed file format. For example, code 920 may be the standard tessellation language (STL) created by 3D Systems' stereolithography CAD program, or the additive manufacturing file (AMF) as an American Society of Mechanical Engineers (ASME) standard, which is an XML-based format designed to allow any CAD software to describe the shape and composition of any three-dimensional object to be manufactured on any AM printer. Code 920 may be transformed between different formats as needed, converted into a set of data signals, and transmitted, received, converted back into code, stored, etc. Code 920 may be an input to system 900 and may come from a part designer, an intellectual property (IP) provider, a design company, an operator or owner of system 900, or other sources. In any case, AM control system 904 executes code 920, dividing object 902 into a series of thin slices to be assembled using AM printer 906 in successive liquid, powder, sheet, or other material layers. In the DMLM example, each layer is melted into the precise geometry defined by code 920 and fused to the previous layer. Subsequently, object 902 may be exposed to any of a variety of finishing processes, such as those described herein for re-shaping or other minor machining, sealing, polishing, etc.

[0055] The terms used herein are for the purpose of describing particular embodiments only and are not intended to limit the present disclosure. As used herein, the singular forms "a", "an", and "the" are intended to include the plural forms as well, unless the context clearly dictates otherwise. It will be further understood that when the terms "comprises" and / or "comprising" are used in the specification, they specify the presence of stated features, integers, steps, operations, elements, and / or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, elements, components, and / or groups thereof.

[0056] This written description uses examples to disclose the invention, including the best mode, and also enables any person skilled in the art to practice the invention, including making and using any device or system and performing any combined method. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. If such other examples have structural elements that are not different from the literal language of the claims, or if they include equivalent structural elements that have no substantial difference from the literal language of the claims, then such other examples are intended to be within the scope of the claims.

Claims

1. A blade, the blade comprising: an internal chamber configured to receive a flow of cooling fluid; a variable thickness wall adjacent to the internal chamber; and a variable thickness impingement plate separating the variable thickness wall from the internal chamber, the impingement plate including a plurality of openings for directing the cooling fluid into an impingement chamber and against the variable thickness wall, wherein the impingement plate is configured to follow the contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to the thickness of the adjacent section of the variable thickness wall; and a plurality of support beams located between the impingement plate and the variable thickness wall, wherein the plurality of support beams are configured to connect the impingement plate to the variable thickness wall, control the separation of the impingement plate from the variable thickness wall, and provide structural support to the variable thickness wall.

2. The blade according to claim 1, wherein the blade comprises a blade of a nozzle in a gas turbine system.

3. The blade according to claim 1, wherein the variable thickness wall comprises one or more of a leading edge wall of the blade, a trailing edge wall of the blade, a pressure side wall of the blade, and a suction side wall of the blade.

4. The blade according to claim 1, wherein the thickness of the variable thickness wall varies according to an expected operating force on the blade.

5. The blade according to claim 4, wherein the thickness of the variable thickness wall is 1.1 times to 2.5 times the nominal wall thickness of the blade.

6. The blade according to claim 1, the blade further comprising a cooling channel formed in a thickened portion of the blade, wherein the cooling channel is fluidly connected to the internal chamber.

7. The blade according to claim 1, the blade further comprising a channel for fluidly connecting the impingement chamber to at least one of an internal chamber of a diaphragm and a trailing edge cooling circuit of the blade.

8. The blade according to claim 7, wherein the blade is integrally formed with the diaphragm.

9. The blade according to claim 1, the blade further comprising at least one channel for fluidly connecting the internal chamber to a wheel space of a gas turbine system.

10. A nozzle segment for a gas turbine system, the nozzle segment comprising: an integrated nozzle and diaphragm, the nozzle including at least one blade, each blade including: an internal chamber configured to receive a flow of cooling fluid; a variable thickness wall adjacent to the internal chamber; and a variable thickness impingement plate separating the variable thickness wall from the internal chamber, the impingement plate including a plurality of openings for directing the cooling fluid into an impingement chamber and against the variable thickness wall, wherein the impingement plate is configured to follow the contour of the variable thickness wall, and wherein the thickness of each section of the impingement plate is inversely proportional to the thickness of the adjacent section of the variable thickness wall; and A plurality of support beams located between the impact plate and the variable-thickness wall, wherein the plurality of support beams are configured to connect the impact plate to the variable-thickness wall, control the separation between the impact plate and the variable-thickness wall, and provide structural support to the variable-thickness wall, and wherein the thickness of the variable-thickness wall varies according to the expected operating force on the blade.

11. The nozzle segment according to claim 10, wherein the variable-thickness wall includes one or more of a leading-edge wall of the blade, a trailing-edge wall of the blade, a pressure side wall of the blade, and a suction side wall of the blade.

12. The nozzle segment according to claim 10, wherein the thickness of the variable-thickness wall is 1.1 times to 2.5 times the nominal wall thickness of the blade.

13. The nozzle segment according to claim 10, the nozzle segment further including a cooling channel formed in a thickened portion of the blade, wherein the cooling channel is fluidly coupled to the internal chamber.

14. The nozzle segment according to claim 10, the nozzle segment further including a channel for fluidly coupling the impact chamber to at least one of an internal chamber of the diaphragm and a trailing-edge cooling circuit of the blade.

15. The nozzle segment according to claim 10, the nozzle segment further including at least one channel for fluidly coupling the internal chamber of the blade to a wheel space of the gas turbine system.

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