Oxidation-activated cooling flow

Abstract

A flow regulation system (200) for increasing the flow rate of a cooling fluid (140, 202) supplied to a cooling system of a component of a gas turbine system (10) is provided. The flow regulation system (200) includes a pneumatic circuit (206) embedded within a section of the component, the pneumatic circuit (206) including a set of interconnected pneumatic passages (208), and a pressure-activated switch (204, 304) fluidly coupled to the pneumatic circuit (206). The pressure-activated switch (204, 304) activates in response to the formation of a break within the section of the component and the exposure of at least one of the pneumatic passages (208) of the pneumatic circuit (206) embedded in the section of the component. Activation of the pressure-activated switch (204, 304) increases the flow rate of the cooling fluid (140, 202) supplied to the cooling system of the component.

Description

[Technical field]

[0001] The present disclosure relates generally to gas turbine systems, and more particularly to a gas turbine system undergoing oxidation. parts providing a secondary flow of cooling fluid to the [Background technology]

[0002] Gas turbine systems are an example of turbomachinery that is widely used in applications such as power generation. Conventional gas turbine systems generally include a compressor section, a combustor section, and a turbine section. During operation of a gas turbine system, turbine blades, nozzle vanes, and Various parts in the system such as shroud segments parts is exposed to a high-temperature gas flow, parts Higher temperature flows generally result in poorer performance, efficiency and overall system life. and Exposed to a hot gas stream to provide increased power output parts to cool the gas turbine system, allowing it to operate at a higher temperature and parts It would be beneficial to extend the life of

[0003] Cooling (e.g., convection cooling, impingement cooling, etc.) is a common feature of gas turbine systems. parts Cooling is often provided by directing a flow of a cooling fluid through internal passages formed in the compressor section of the gas turbine system. Often the cooling fluid is provided by bleeding a portion of the air discharged by the compressor section of the gas turbine system.

[0004] Thermal Barrier Coating (TBC) provides a protective heat shield to prevent damage from high temperatures and oxidation and By reducing thermal fatigue parts They are often applied to gas turbine system components to extend their life. Spallation of TBCs is a common problem in gas turbine systems. When a TBC spalls, a portion of the TBC breaks off and partsThis can cause the material to separate from the underlying surface, exposing it to high temperatures and potentially damaging it (e.g., by oxidation). Summary of the Invention

[0005] The first embodiment is a gas turbine system parts 1. A flow regulation system for increasing a flow rate of a cooling fluid supplied to a cooling system of the present invention, comprising: parts a pneumatic circuit embedded within a section of the pneumatic circuit, the pneumatic circuit including a set of interconnected pneumatic passages; and a pressure actuated switch fluidly coupled to the pneumatic circuit, the pressure actuated switch including: parts In the above section crack Formation and parts At least one of the pneumatic passages of the pneumatic circuit embedded within said section of Air pressure passage and a pressure activated switch that is activated in response to exposure of the pressure activated switch to the pressure activated switch. parts The present invention relates to a flow regulation system for increasing the flow rate of cooling fluid supplied to a cooling system of a vehicle.

[0006] Another embodiment is a system for controlling a gas turbine parts of crack 1. A method for controlling oxidation in the vicinity of parts monitoring pressure within a pneumatic circuit embedded within a section of the pneumatic circuit, the pneumatic circuit including a set of interconnected pneumatic passages; and in response to a decrease in pressure within the pneumatic circuit, parts Auxiliary flow of cooling fluid to the cooling system of or Providing an increased flow rate of cooling fluid, the auxiliary flow of cooling fluid being or The increased flow rate of cooling fluid parts of crack Oxidation in the vicinity of Prevention or The present invention provides a method including the steps of: preventing; and providing.

[0007] A further embodiment is a system comprising a turbine shroud, a cooling system for cooling the turbine shroud, and a flow regulation system for increasing a flow rate of a cooling fluid provided to the cooling system of the turbine shroud, the flow regulation system being a pneumatic circuit embedded within a section of the turbine shroud, the pneumatic circuit including a set of interconnected pneumatic passages, and a pressure actuated switch fluidly coupled to the pneumatic circuit, the pressure actuated switch controlling a flow rate of the cooling fluid provided to the cooling system of the turbine shroud. crack Formation and At least one of the pneumatic passages of the pneumatic circuit embedded within said section of the turbine shroud. Air pressure passage and a pressure activated switch actuated in response to exposure of the pressure activated switch, wherein actuation of the pressure activated switch increases a flow of cooling fluid supplied to a cooling system of the turbine shroud.

[0008] Exemplary aspects of the present disclosure address the problems described herein. and / or Resolve other issues not discussed.

[0009] These features of the present disclosure and Other features will be more readily understood from the following detailed description of the various aspects of the disclosure when considered in conjunction with the accompanying drawings which illustrate various embodiments of the disclosure. [Brief description of the drawings]

[0010] [Figure 1] 1 is a schematic diagram of a gas turbine system according to an embodiment. [Diagram 2] FIG. 1 illustrates a side view of a portion of a turbine section of a gas turbine system according to an embodiment. [Diagram 3] FIG. 2 is an isometric view of a turbine shroud according to an embodiment. [Figure 4] FIG. 4 is a cross-sectional view of the turbine shroud of FIG. 3 according to an embodiment. [Diagram 5] FIG. 1 illustrates a pressure activated switch in an unactuated state according to an embodiment. [Figure 6] FIG. 1 illustrates a pressure activated switch in an activated state according to an embodiment. [Figure 7] FIG. 13 illustrates a pressure activated switch in an unactuated state according to another embodiment. [Figure 8] FIG. 13 illustrates a pressure activated switch in an activated state according to another embodiment. [Figure 9] FIG. 1 is a block diagram of an additive manufacturing process including a non-transitory computer readable storage medium storing code representing a turbine shroud according to an embodiment of the disclosure. [Figure 10] 1 is a flow diagram of a method for preventing and / or inhibiting oxidation according to an embodiment of the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0011] It should be noted that the drawings of the present disclosure are not necessarily to scale. The drawings are intended to illustrate only typical aspects of the present disclosure, and therefore should not be considered as limiting the scope of the present disclosure. In the drawings, like numbers represent like elements between the drawings.

[0012] Reference will now be made in detail to the exemplary embodiments, as illustrated in the accompanying drawings. It should be understood that the following description is not intended to limit the embodiments to a single preferred embodiment. To the contrary, the spirit of the described embodiments, as defined by the appended claims, is intended to be illustrative, not restrictive, and is not intended to be limiting. and Alternatives, modifications, and variations that may fall within the scope of the present invention. and It is intended to cover equivalent examples.

[0013] As a first issue, in order to clearly explain the present disclosure, the relevant machines within the scope of this disclosure parts When describing the subject matter with reference to the specification, it becomes necessary to select specific terminology. When doing this, whenever possible, common industry terms are used with their accepted meaning. and Unless otherwise indicated, such terminology has the same meaning as in the context of this application. andThe appended claims should be given a broad interpretation consistent with their scope. parts There are several different or It will be understood that overlapping terminology may be used to refer to what may be described herein as a single component. parts may contain and be referenced in other contexts as consisting of parts Content described herein as comprising may be referred to elsewhere as a single component.

[0014] In addition, several descriptive terms may be used periodically herein and it will prove useful to define these terms at the outset here. These terms and The definitions are as follows, unless otherwise stated: As used herein, "downstream" and "Upstream" refers to the working fluid passing through the turbine; or For example, the flow of air through a combustor or a turbine parts A term that indicates a direction relative to the flow of a fluid, such as a coolant, through one of the systems. The term "downstream" corresponds to the direction of fluid flow, while the term "upstream" refers to the opposite direction of flow. and The term "aft" refers to the direction, unless otherwise specified, and "forward" refers to the front of the engine. or Refers to the compressor end, and "rear" refers to the rear of the engine or Refers to the turbine end. In addition, "leading" and The terms "following" and "forward" are interchangeable. and Used in a similar description to the term "backward" and / or In many cases, different radial, axial, and and / or It becomes necessary to account for the parts in their circumferential positions. The "A" axis represents an axial orientation. As used herein, "axial" and / orThe term "axially" refers to the relative position / orientation of an object along an axis A that is substantially parallel to the axis of rotation of the gas turbine system (particularly the rotor section). As used further herein, "radially" refers to the relative position / orientation of an object along an axis A that is substantially parallel to the axis of rotation of the gas turbine system (particularly the rotor section). and / or The term "radially" refers to the relative position / orientation of an object along a direction "R" (see FIG. 1) that is substantially perpendicular to and intersects axis A at only one location. Finally, the term "circumferential" refers to the movement around axis A. or Refers to a location (e.g., direction "C").

[0015] In various embodiments, or Described as "in fluid communication" parts is one or The bonding can occur along multiple interfaces. In some embodiments, these interfaces can be separate. parts In other cases, these interfaces may include joints between and / or It may include interconnects that are integrally formed, i.e., in some cases, "bonded" to one another. parts may be formed simultaneously to define a single continuous member. However, in other embodiments, these joined parts can be formed as separate pieces and then joined by known processes (eg, fastening, ultrasonic welding, adhesives).

[0016] An element or When a layer is referred to as being "on," "engaged," "connected," or "coupled" to another element, it may be directly on, engaged, connected, or coupled to the other element, or there may be intervening elements. In contrast, when an element is referred to as being "directly on," "directly engaged," "directly connected," or "directly coupled" to another element, there may be no intervening elements. orThere may be no layers. Other words used to describe relationships between elements should be interpreted similarly (e.g., "between" vs. "directly between," "adjacent to" vs. "directly adjacent to," etc.). As used herein, " and / or The term " " refers to one of the associated enumerated items. or Includes any and all combinations of multiples.

[0017] FIG. 1 illustrates a schematic diagram of a gas turbine system 10 according to various embodiments. As shown, the gas turbine system 10 includes a compressor section 12 for compressing an incoming airflow 14 and delivering a flow of compressed air 16 to a combustor section 18. The combustor section 18 mixes the flow of compressed air 16 with a pressurized supply of fuel 20 and ignites the mixture to generate a flow of combustion gases 22. Although only a single combustor section 18 is shown, the gas turbine system 10 may include any number of combustor sections 18. The flow of combustion gases 22 is then delivered to a turbine section 24. The flow of combustion gases 22 drives the turbine section 24 to generate mechanical work. The mechanical work generated in the turbine section 24 may drive the compressor section 12 via a shaft 26 and may further be used to drive an external load 28, such as an electrical generator.

[0018] FIG. 2 illustrates at least one stage 30 (one shown) of turbine blades 32. and Within the casing 38 of the turbine section 24 ("turbine casing 38"). Placed 2 illustrates a side view of a portion of the turbine section 24 of a gas turbine system including at least one stage 34 (one shown) of nozzles 36 arranged in a radially extending manner about the rotor 26. Each stage 30 of turbine blades 32 is coupled to the rotor 26 and is circumferentially disposed about the rotor 26. Placed 2. Each stage 34 of nozzles 36 is coupled to a turbine casing 38 of the turbine section 24 and extends circumferentially therearound. PlacedThe nozzles 36 are arranged in a circular arc.

[0019] In the embodiment shown in FIG. 2, each nozzle 36 is disposed between an outer platform 42 and an inner platform 44. Placed Similar to the nozzle 36, each turbine blade 32 in the turbine section 24 includes an airfoil 46 that extends radially from the rotor 26. Each airfoil 46 includes a tip portion 48 and a distal end portion 49 opposite the tip portion 48. Placed and a platform 50.

[0020] Turbine Blade 32 and The nozzles 36 are arranged axially adjacent to one another within a turbine casing 38. Place In FIG. 2, for example, adjacent to the turbine blade 32 in the axial direction and downstream of the turbine blade 32, Placed The turbine section 24 is shown axially extending through a turbine casing 38. Placed A plurality of stages 30 of interdigitated turbine blades 32 and Multiple stages 34 of nozzles 36 may be included.

[0021] The turbine section 24 of the gas turbine system 10 is axially aligned through a turbine casing 38. Placed 2, for example, a stage 52 of shrouds 54 may be radially adjacent to a stage 30 of turbine blades 32. Placed This essentially encloses or The stage 52 of the shroud 54 is also axially adjacent the stage 34 of the nozzle 36. and / or Upstream Place Further, the stage 52 of the shroud 54 may be disposed between two adjacent stages 34 of the nozzles 36 located on opposite sides of the stage 30 of the turbine blades 32. PlaceThe stages 52 of the shroud 54 may be coupled about the turbine casing 38 using a set of extensions 56 that each include an opening 58 configured to receive a corresponding section of the shroud 54.

[0022] An isometric view of the turbine shroud 54 is shown in Figure 3, and a cross-sectional view of the turbine shroud 54 is shown in Figure 4. As shown, the turbine shroud 54 includes a body 60. The body 60 of the turbine shroud 54, as well as various other components of the turbine shroud 54, are Parts and / or The features may be formed using any suitable technique, including additive manufacturing processes. For example, the turbine shroud 54, including the body 60, may be formed using a variety of techniques, including direct metal laser melting (DMLM) (also referred to as selective laser melting (SLM)), direct metal laser sintering (DMLS), electron beam melting (EBM), stereolithography (SLA), binder jetting, and the like. or It may also be formed by any other suitable additive manufacturing process.

[0023] The body 60 of the turbine shroud 54 includes a support portion 62, an intermediate portion 64, and a seal portion 66. The support portion 62 is directly coupled to the turbine shroud 54 and / or or Turbine casing 38 and 2. The support portion 62 includes a forward end 68 including at least one forward hook 70, an aft end 72 including at least one aft hook 74, a first surface 76, and a second surface 78. The intermediate portion 64 includes a non-linear segment 84. and Various features of the body 60 between the opposing slash faces 80, 82 of the body 60 include a forward segment 86, for example, which forms a seal within the turbine section 24 and defines a hot gas flow path of the combustion gases 22 flowing through the turbine section 24. and / orThe seal portion 66 may be adapted to at least partially define a flowpath of the combustion gases 22 flowing through the turbine section 24. The seal portion 66 may be adapted to be adjacent a hot gas flowpath of the combustion gases 22 in the turbine section 24. Place The hot gas path (HGP) surface 88 may include

[0024] The body 60 of the turbine shroud 54 is supported within a first surface 76 of the support portion 62 between a forward end 68 and an aft end 72 of the body 60 . and / or The body 60 further includes at least one inlet opening 90 formed therethrough. The inlet opening 90 is formed through the support portion 62, the intermediate portion 64, and formed through the seal portion 66, and / or It is in fluid communication with a cooling circuit 92 (FIG. 4) contained therein.

[0025] The turbine shroud 54 may also include a set of metering plates 94 (shown in dashed lines in FIG. 3 ) coupled to the first surface 76 of the support portion 62 of the body 60. Each metering plate 94 (only one shown) overlies a respective inlet opening 90 and and / or The metering plate 94 may be attached to the first surface 76 so as to at least partially cover it. The metering plate 94 allows a predetermined flow rate of cooling fluid to enter the cooling circuit 92 via the inlet opening 90.

[0026] Various plenum(s) of the turbine shroud 54 and / or The cooling passage(s) are shown in Figure 4, which is a cross-sectional view of the turbine shroud 54 of Figure 3. As shown, the turbine shroud 54 includes a cooling passage formed through a portion of the body 60 of the turbine shroud 54. and / or The turbine shroud 54 includes at least one plenum 100 that may extend through the support portion 62, the intermediate portion 64, and the main body 60 of the turbine shroud 54. andThe intermediate portion 64 of the body 60 may extend radially through at least a portion of the sealing portion 66. and One of the plenums 100 formed within the seal portion 66 or The plurality of portions are disposed between the opposing slash faces 80, 82. and / or Although only a single plenum 100 is shown, it is understood that the turbine shroud 54 may include additional plenums.

[0027] The plenum 100 is fluidly coupled to the inlet opening(s) 90 formed in the support portion 62 of the body 60. and / or As described herein, the plenum 100 may be configured to receive a supply of a cooling fluid 140 (e.g., compressor discharge air) via the inlet opening(s) 90 to provide the cooling fluid 140 to separate cooling passages formed in the turbine shroud 54 for cooling the turbine shroud 54 during operation of the gas turbine system 10.

[0028] As shown in FIG. 4, the turbine shroud 54 is formed, positioned, and secured within a body 60 of the turbine shroud 54. and / or The cooling passage 102 extends through the plenum 100 and includes a first cooling passage 102 in fluid communication with the plenum 100. More specifically, the first cooling passage 102 extends between a forward end 104 and an aft end 106 of the body 60. and / or Adjacently, within the seal portion 66 of the body 60 of the turbine shroud 54 Placed R, and / or In addition, the first cooling passage 102 may extend between the opposing slashfaces 80, 82. and / or Adjacent thereto, the first cooling passage 102 may extend through the seal portion 66 of the body 60. The first cooling passage 102 may also extend radially within the seal portion 66 between the plenum 100 and the HGP surface 88 of the seal portion 66. Placed It is possible.

[0029] The first cooling passage 102 is made up of a number of separate segments, sections, and / or For example, the first cooling passage 102 may be positioned between a forward section 110 and an aft section 112, and / or 4, the central portion 108 of the first cooling passage 102 is formed centrally between the forward end 104 and the aft end 106 of the seal portion 66. and / Or placed The forward portion 110 of the first cooling passage 102 is immediately adjacent the forward end 104 of the seal portion 66 and axially adjacent the central portion 108. and / or Formed axially upstream and / Or placed Similarly, the aft portion 112 of the first cooling passage 102 is formed opposite the forward portion 110 and directly adjacent the aft end 106 of the seal portion 66. and / Or placed The central portion 108 is adapted to improve the operating efficiency of the turbine section 24. and / or To improve the operational life of the turbine shroud 54 in the turbine section 24, the areas in the turbine shroud 54 most requiring cooling and / or or It may be radially aligned with the axial portion of the HGP surface 88 of the seal portion 66 that requires maximum heat exchange.

[0030] Plenum 100 and The first cooling passages 102 are separated by ribs 114. The ribs 114 may extend within the body 60 of the turbine shroud 54 between the opposing slashfaces 80, 82. A plurality of impingement openings 116 are provided through the ribs 114. The impingement openings 116 fluidly couple the plenum 100 and the central portion 108 of the first cooling passages 102. During operation of the gas turbine system 10, a cooling fluid 140 flows from the plenum 100 through the plurality of impingement openings 116 to the central portion 108 of the first cooling passages 102.

[0031] In addition to the first cooling passage 102, the body 60 of the turbine shroud 54 also includes a second cooling passage 118. As shown in FIG. 4, the second cooling passage 118 may extend into the body 60 of the turbine shroud 54 adjacent the forward end 104 of the seal portion 66 and may extend into the seal portion 66 between the opposing slashfaces 80, 82. The second cooling passage 118 may be adjacent to and upstream of the central portion 108 of the first cooling passage 102. Placed 1, radially inward from the forward portion 110 of the first cooling passage 102. Placed Additionally, a second cooling passage 118 is formed between the forward portion 110 of the first cooling passage 102 and the HGP surface 88 of the seal portion 66. Or placed can be.

[0032] A rib 120 separates the second cooling passage 118 from the forward portion 110 of the first cooling passage 102. The rib 120 may extend within the body 60 of the turbine shroud 54 between the opposing slashfaces 80, 82. The second cooling passage 118 is in direct fluid communication with the forward portion 110 of the first cooling passage 102. A plurality of impingement openings 122 formed through the rib 120 fluidly couple the forward portion 110 of the first cooling passage 102 and the second cooling passage 118. During operation of the gas turbine system 10 (see FIG. 1 ), cooling fluid 140 flowing through the forward portion 110 of the first cooling passage 102 passes from the forward portion 110 of the first cooling passage 102 through the impingement openings 122. or It flows to the second cooling passage 118 .

[0033] The body 60 of the turbine shroud 54 further includes a plurality of forward exhaust holes 124 in flow communication with the second cooling passage 118. During operation, the plurality of forward exhaust holes 124 discharge cooling fluid 140 from the second cooling passage 118 into the hot gas path of the combustion gases 22 flowing through the turbine section 24 of the gas turbine system 10.

[0034] A third cooling passage 126 is provided within the seal portion 66 of the body 60 of the turbine shroud 54. As shown in FIG. 4, the third cooling passage 126 is disposed adjacent the aft end 106 of the seal portion 66. The first cooling passage 102 and Similar to the second cooling passage 118, the third cooling passage 126 may also extend within the seal portion 66 between the opposing slashfaces 80, 82. The third cooling passage 126 is adjacent to and downstream of the central portion 108 of the first cooling passage 102. Placed 1 and 2, extending radially inwardly from the rear portion 112 of the first cooling passage 102 adjacent the HGP surface 88 of the seal portion 66. Placed can be.

[0035] A rib 128 separates the third cooling passage 126 from the aft portion 112 of the first cooling passage 102. The rib 128 may extend within the body 60 of the turbine shroud 54 between the opposing slashfaces 80, 82. A plurality of impingement openings 130 formed through the rib 128 fluidly couple the aft portion 112 of the first cooling passage 102 and the third cooling passage 126. During operation of the gas turbine system 10 (see FIG. 1 ), cooling fluid 140 flowing through the aft portion 112 of the first cooling passage 102 passes from the aft portion 112 of the first cooling passage 102 through the impingement openings 130. or It flows to the third cooling passage 126 .

[0036] The body 60 of the turbine shroud 54 further includes a plurality of aft exhaust holes 132 in flow communication with the third cooling passage 126. During operation, the plurality of aft exhaust holes 132 are configured to discharge cooling fluid 140 from the third cooling passage 126 into a hot gas path of the combustion gases 22 flowing through the turbine section 24 of the gas turbine system 10.

[0037] During operation of the gas turbine system 10, a cooling fluid 140 (e.g., compressor discharge air) flows under pressure through the body 60 to cool the turbine shroud 54. More specifically, as the turbine shroud 54 is exposed to the combustion gases 22 flowing through the hot gas path of the turbine section 24 during operation of the gas turbine system 10, the cooling fluid 140 is provided to various cooling mechanisms (e.g., plenum 100, cooling passages 102, 118, 126, exhaust channels 124, 132, etc.) within the body 60 to cool the turbine shroud 54. In a non-limiting example, the cooling fluid 140 enters the plenum 100 of the body 60 of the turbine shroud 54 through inlet opening(s) 90 formed in the first surface 76 of the support portion 62 of the body 60 of the turbine shroud 54. Additionally, the turbine shroud 54 is provided with a cooling passage 90 formed in the first surface 76 of the support portion 62 of the body 60 of the turbine shroud 54. and / or When including metering plate(s) 94 attached to first surface 76 at least partially covering it, metering plate(s) 90 may regulate a fixed flow amount of cooling fluid 140 through inlet opening(s) 90 into plenum(s) 100. and Adjust the pressure.

[0038] The cooling fluid 140 is configured to flow radially from the inlet opening(s) 90 through the plenum 100 toward the cooling passages 102, 118, 126 formed in the seal portion 66 of the body 60 of the turbine shroud 54. More specifically, the cooling fluid 140 provided to the plenum 100 flows radially through the plenum 100 toward the ribs 114 and then through the impingement openings 116 in the ribs 114 to the central portion 108 of the first cooling passage 102. From the central portion 108 of the first cooling passage 102, the cooling fluid 140 flows axially toward the forward portion 110 of the first cooling passage 102. and It flows into the rear section 112 .

[0039] A portion of the cooling fluid 140 entering the forward portion 110 of the first cooling passage 102 enters the second cooling passage 118 through impingement openings 122 formed in the ribs 120. Similarly, a portion of the cooling fluid 140 entering the aft portion 112 of the first cooling passage 102 enters the third cooling passage 126 through impingement openings 130 formed in the ribs 128.

[0040] From the second cooling passage 118, a portion of the cooling fluid 140 may exit the body 60 of the turbine shroud 54 through the exhaust holes 124. Additionally, a portion of the cooling fluid 140 in the third cooling passage 126 may exit the body 60 through the exhaust holes 132. The remaining cooling fluid 140 that is not exhausted through the exhaust holes 124, 132 may be supplied to other cooling features within the body 60 of the turbine shroud 54.

[0041] within the HGP surface 88 of the turbine shroud 54 as a result of oxidation after TBC spallation. crack may occur. crack teeth, crack Continued oxidation of the HGP surface 88 in the vicinity of the turbine shroud 54 can increase in size and shorten the operating life of the turbine shroud 54.

[0042] According to an embodiment, the HGP surface 88 crack In response to the formation of the cooling fluid 202 within the HGP surface 88, a flow regulation system 200 is provided for delivering a secondary flow of cooling fluid 202 to the cooling circuit 92 of the turbine shroud 54. crack In response to the formation of the secondary flow of cooling fluid 202, the flow regulation system 200 is configured to direct a secondary flow of cooling fluid 202 into the cooling circuit 92 (e.g., into the plenum 100) within the body 60 of the turbine shroud 54. The flow regulation system 200 may be configured to, for example, direct a secondary flow of cooling fluid 202 into the cooling circuit 92 within the body 60 of the turbine shroud 54, as shown in FIG. and6, the flow regulation system 200 may be used to provide a supplemental flow of cooling fluid 202 to the cooling circuit 92 of the turbine shroud 54 having the cooling fluid 140 supplied directly from a discharge (e.g., first stage or higher) of the compressor section 12 (FIG. 1) of the gas turbine system 10. According to another embodiment, the flow regulation system 200 may be configured as shown in FIG. and 8, the flow regulation system 200 may be used to increase the flow of cooling fluid 140 to the cooling circuit 92 of a turbine shroud 54 that receives its cooling fluid 140 through an extraction port (e.g., second stage or higher) of the compressor section 12 of the gas turbine system 10. Although described herein with respect to the turbine shroud 54, the flow regulation system 200 may also be used with other components of the gas turbine system 10 that may be subject to oxidation. parts For example, the flow regulation system 200 may be configured for use with parts On the HGP surface crack In response to the formation of parts Auxiliary supply of cooling fluid to the cooling system ( or The cooling fluid may be configured to provide an increased flow rate of the cooling fluid.

[0043] According to various embodiments, the flow regulation system 200 includes a pressure activated switch 204 (FIG. 5 and Figure 6) or 304 (Figure 7 and 8), and a pneumatic circuit 206 (FIG. 3) that includes a set of interconnected pneumatic passages 208 that are completely embedded within the HGP surface 88. and At least one of the air pressure passages 208 is fluidly coupled to a pressure activated switch 204, 304. and 4, the air pressure passages 208 may be arranged in a grid pattern within the HGP surface 88. Many other configurations may be used. For example, the air pressure passages 208 may be provided in a sinusoidal configuration, a mesh configuration, a spiral configuration, a rectangular grid, etc. In general, the arrangement of the air pressure passages 208 within the HGP surface 88 may be and The spacing is relatively small for HGP Surface 88 crackEven if one of the air pressure passages 208 in the HGP surface 88 or It's something that gives you multiple exposures. crack , none of the pneumatic passages 208 of the pneumatic circuit 206 are exposed.

[0044] During normal operation (e.g., HGP Surface 88 crack In the absence of a pressure switch 204, 304, the pressure in the pneumatic passage 208 of the pneumatic circuit 206 is sufficient to maintain the pressure activated switch 204, 304 in an inactivated state. crack , one of the pneumatic passages 208 of the pneumatic circuit 206 embedded within the HGP surface 88 or When exposed, pressure drops within the pneumatic passage 208 of the pneumatic circuit 206, resulting in actuation of the pressure activated switch 204, 304. When actuated, the pressure activated switch 204, 304 closes the auxiliary supply ( or 1. The increased flow rate of cooling fluid 140 into the cooling system 92 of the turbine shroud 54. Advantageously, the increased flow rate of cooling fluid into the cooling system 92 allows: crack Is oxidation of the HGP surface 88 in the vicinity of the or It's like being blocked.

[0045] An embodiment of a pressure-activated switch 204 is shown in FIG. and As shown in FIG. 6, the pressure actuated switch 204 is crack 5, for example, the fluid inlet 212 of the pressure actuated switch 204 is fluidly coupled to a compressor discharge chamber (CDC), which receives a flow of compressed air 16 ( FIG. 1 ) from the compressor portion 12 of the gas turbine system 10. The pressure actuated switch 204 is configured to selectively provide a supplemental supply of cooling fluid 202 to the cooling system 92 of the turbine shroud 54 in response to the formation of a pressure actuated switch 204 pressure outlet 210 and a fluid inlet 212 fluidly coupled to a pressurized source of cooling fluid. In FIG. 5, for example, the fluid inlet 212 of the pressure actuated switch 204 is fluidly coupled to a compressor discharge chamber (CDC), which receives a flow of compressed air 16 ( FIG. 1 ) from the compressor portion 12 of the gas turbine system 10.

[0046] The pressure activated switch 204 further includes a piston 214 and a disk 216 coupled to the piston 214 that is configured to mate with a seat 218. A biasing element 220, such as a spring, biases the disk 216 toward the seat 218. The distal end of the piston 214 is connected to one of the pneumatic passages 208. or The pneumatic circuit 206 is fluidly coupled via multiple

[0047] HGP Surface 88 crack 5, the pressure in the air pressure passage 208 with the biasing force applied by the biasing element 220 is greater than the pressure in the compressor discharge chamber. Thus, the disk 216 coupled to the piston 214 is forced against the seat 218, closing the pressure activated switch 204. This prevents the auxiliary supply of cooling fluid 202 from flowing from the compressor discharge chamber to the cooling system 92 of the turbine shroud 54.

[0048] The pressure actuated switch 204 is crack Formation of (e.g., due to oxidation) and 2. The pressure actuated switch 204 is operative in response to exposure of at least a portion of the pneumatic passages 208 of the pneumatic circuit 206 embedded within the HGP surface 88. The exposure causes a loss of pressure within the pneumatic passages 208 of the pneumatic circuit 206. As a result, as shown in FIG. 6, the pressure within the compressor discharge chamber becomes greater than the biasing force applied by the biasing element 220 at this point, forcing the disk 216 coupled to the piston 214 away from the seat 218 and fluidly coupling the fluid inlet 212 to the fluid outlet 210 of the pressure actuated switch 204. A supplemental supply of cooling fluid 202 may now flow from the compressor discharge chamber through the pressure actuated switch 204 to the cooling system 92 of the turbine shroud 54. The increased flow rate of cooling fluid to the cooling system 92 may crack Prevents oxidation of the HGP surface 88 in the vicinity of or Block.

[0049] Another embodiment of a pressure activated switch 304 is shown in FIG. andAs shown in FIG. 8, the pressure actuated switch 304 is crack 3. The pressure activated switch 304 is configured to selectively trigger an increase in the flow rate of the cooling fluid 140 provided to the cooling system 92 of the turbine shroud 54 in response to the formation of a pressure drop across the turbine shroud 54. The pressure activated switch 304 may be used, for example, in the turbine shroud 52 that receives its cooling fluid 140 via an extraction port (e.g., second stage or higher) of the compressor section 12 of the gas turbine system 10.

[0050] The pressure actuated switch 304 is mounted on the body 60 of the turbine shroud 54. Place The pressure actuated switch 304 may be connected to one of the pneumatic passages 208 of the pneumatic circuit 206 via a port 308. or The deformable chamber 306 includes a plurality of fluidly coupled deformable chambers 306. The port 308 is disposed on a first surface of the deformable chamber 306. Placed The electrical current passes through a first electrical contact 310 disposed on a second opposing surface 314 of the deformable chamber 306. A second electrical contact 312 is provided on a second opposing surface 314 of the deformable chamber 306. The first surface 314 of the deformable chamber 306 is exposed to the pressure in a compressor discharge chamber, which receives a flow of compressed air 16 (FIG. 1) from the compressor portion 12 of the gas turbine system 10.

[0051] HGP Surface 88 crack 7, in the air pressure passage 208 ( and The pressure in the deformable chamber 306 is higher than the pressure in the compressor discharge chamber. The higher pressure in the deformable chamber 306 causes the deformable chamber 306 to expand, preventing the second electrical contact 312 from contacting the first electrical contact 310. Thus, the pressure activated switch 304 is in an unactuated state. The first electrical contact 310 and The second electrical contact 312 is electrically connected to a controller 316 .

[0052] The pressure actuated switch 304 is crack Formation (e.g., due to oxidation) and8, the pressure activated switch 304 is actuated in response to exposure of at least a portion of the pneumatic passage 208 of the pneumatic circuit 206 embedded within the HGP surface 88. The exposure causes a loss of pressure within the pneumatic passage 208 of the pneumatic circuit 206. As a result, the pressure in the compressor discharge chamber is greater than the pressure in the deformable chamber 306 at this point. As shown in FIG. 8, this causes the deformable chamber 306 to collapse. As a result, the first electrical contact 310 is displaced toward and against the second electrical contact 312, completing an electrical circuit and actuating the pressure activated switch 304.

[0053] In response to actuation of the pressure activated switch 304, the controller 316 increases the amount of cooling fluid 140 flowing through the extraction port 318 into the cooling system 92 of the turbine shroud 54. The increased flow rate of cooling fluid 140 may be provided, for example, by increasing the size of the opening of the extraction port 318, as shown in dashed lines in FIG. 8, in response to a signal from the controller 316. The increased flow rate of cooling fluid 140 into the cooling system 92 may be provided, for example, by increasing the size of the opening of the extraction port 318, as shown in dashed lines in FIG. crack Prevents oxidation of the HGP surface 88 in the vicinity of or Block.

[0054] Oxidation according to the embodiment Prevention and / or A flow diagram of a method for blocking is shown in Figure 10 and will be described with reference to Figures 3-8. At S1, pressure in the pneumatic passage 208 of the pneumatic circuit 206 is monitored by the flow monitoring system 200. In response to a drop in pressure in the pneumatic circuit 206 (YES at S2), flow proceeds to S3. In response to no drop in pressure in the pneumatic circuit 206 (NO at S2), flow returns to S1.

[0055] At S3, the pressure activated switch 204, 304 is activated in response to a drop in pressure in the pneumatic circuit 206. At S4, the auxiliary flow of the cooling fluid 202 is or An increased flow rate of cooling fluid 140 is provided to the cooling system 92 of the turbine shroud 54. At S5, the HGP surface 88 of the turbine shroud 54 crackThe oxidation in the vicinity of the cooling fluid 202 is or Is it hindered due to an increase in the flow rate of the cooling fluid 140? or Be prevented.

[0056] As discussed above, the body 60 of the turbine shroud 54, as well as the various components of the flow regulation system 200 disclosed herein, parts / Various other parts of turbine shroud 54 including Parts and / or The features may be formed using any suitable technique, including an additive manufacturing process, which may use any suitable material capable of withstanding the operating characteristics (e.g., exposed temperatures, exposed pressures, etc.) experienced by the turbine shroud 54 in the gas turbine system 10 during operation.

[0057] As used herein, additive manufacturing (AM) may include any process that produces an object through the creation of successive layers of material, rather than the removal of material as in traditional processes. Additive manufacturing involves the use of any type of tool, mold, or other or No tools are used and very little waste material is produced or Complex geometric shapes can be formed without any plastic, much of which would otherwise be cut away and discarded. or From a solid billet of metal parts Instead of machining a part, the material used in additive manufacturing is only the material needed to produce the part. Additive manufacturing processes 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) can be included. In the current setup, DMLM or SLM has proven to be advantageous.

[0058] To illustrate an example of an additive manufacturing process, FIG. 9 shows a schematic / block diagram of an exemplary computerized additive manufacturing system 900 for generating an object 902. In this example, the system 900 is configured for DMLM. It should be understood that the general teachings of this disclosure are applicable to other forms of additive manufacturing as well. The object 902 is shown as a turbine shroud 54 (see FIGS. 2-8). The AM system 900 generally includes a computerized additive manufacturing (AM) control system 904 and an AM printer 906. The AM system 900 executes code 920 including a set of computer executable instructions defining the turbine shroud 54 to physically generate the object 902 using the AM printer 906, as described. Each AM process may use various feedstock materials, for example in the form of fine powder, liquid (e.g., polymer), sheets, etc., the stock of which may be held in a chamber 910 of the AM printer 906. In this case, the turbine shroud 54 is made of a metal capable of withstanding the environment of the gas turbine system 10 (see FIG. 1). or As shown, the applicator 912 may create a thin layer of raw material 914 that is spread as a blank canvas on the build plate 915 of the AM printer 906, from which each successive slice of the final object is created. In other cases, the applicator 912 may apply the next layer directly onto the previous layer as defined by code 920, for example when a metal binder jetting process is used. or It may also be printed. In the example shown, it is laser or The electron beam 916 fuses the grains of each slice together, as defined by code 920, although this may not be necessary if a fast-setting liquid plastic / polymer is employed. Various parts of the AM printer 906 may move to accommodate the addition of each new layer, for example, after each layer the build platform 918 may lower, and / or Chamber 910 and / or The applicator 912 may rise.

[0059] The AM control system 904 is shown implemented as computer program code on a computer 930. In this regard, the computer 930 is shown to include a memory 932, a processor 934, an input / output (I / O) interface 936, and a bus 938. Additionally, the computer 930 may include external I / O devices / resources 940. and The processor 934 is shown in communication with the memory 932 under instructions from the code 920 representing the turbine shroud 54 described herein. and / or The processor 934 executes computer program code, such as the AM control system 904, stored in the storage system 942. When executing the computer program code, the processor 934 communicates with the memory 932, the storage system 942, the I / O devices 940, and / or Reading data from the AM printer 906 and / or Data can be written to these. parts The I / O devices 940 may include any device 940 that allows a user to interact with the computer (e.g., a keyboard, a pointing device, a display, etc.). The computer 930 includes hardware and The software is merely representative of various possible combinations. For example, the processor 934 may include a single processing unit, or One or Multiple locations, e.g. clients and One on the server or The memory 932 may be distributed across multiple processing units. and / or The storage system 942 includes one or The memory 932 may be located in multiple physical locations. and / orThe storage system 942 may comprise 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 may comprise any type of computing device, such as a network server, a desktop computer, a laptop, a mobile device, a mobile phone, a pager, a personal digital assistant, etc.

[0060] The additive manufacturing process begins by storing code 920 representing turbine shroud 54 in a non-transitory computer-readable storage medium (e.g., memory 932, storage system 942, etc.). For example, code 920 may include a precisely defined 3D model of turbine shroud 54 and may be generated from any of a variety of well-known computer-aided design (CAD) software systems, such as AutoCAD®, TurboCAD®, DesignCAD 3D Max, etc. In this regard, code 920 may be generated from any currently known 3D model of turbine shroud 54. or It may be a file format developed later. For example, Code 920 is a file format created for the Standard Tessellation Language (STL) for 3D Systems' stereolithography CAD program. or The AM printer may be configured to generate an AM file (Additive Manufacturing File) that is an American Society of Mechanical Engineers (ASME) standard that is an Executable Markup Language (XML) based format that allows any CAD software to generate the geometry of any three-dimensional object that is to be manufactured by the AM printer. and The code 920 is designed to be able to represent a variety of configurations, and can be used to convert between different formats and convert into a set of data signals, if necessary. and Transmit and receive as a set of data signals and The code 920 may be an input to the system 900 and may be provided by a part designer, an intellectual property (IP) provider, a design firm, an operator or owner of the system 900, orIn any case, the AM control system 904 executes the code 920 to divide the turbine shroud 54 into successive thin slices, which may be formed by any of a variety of methods, including liquid, powder, sheet, or other sources. or Successive layers of other materials are assembled using an AM printer 906. In the DMLM example, each layer is melted and fused to the previous layer to a precise geometric shape defined by code 920. The turbine shroud 54 can then be subjected to any of a variety of finishing processes, such as recontouring. or It may undergo the finishing processes described herein for other minor machining, sealing, polishing, etc.

[0061] The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the disclosure. As used herein, the singular forms "a," "an," "an and "The" is intended to include the plural unless expressly stated otherwise. and / or The term "comprising" as used herein means any combination of stated features, integers, steps, operations, elements, and / Or parts indicates that there is one or A plurality of other features, integers, steps, operations, elements, parts , and / or The existence of those pairs or It will further be understood that addition is not excluded.

[0062] This written description is provided to disclose the invention, including the best mode thereof, and Any person skilled in the art would be able to or Preparation of the system andThe examples are used to enable the practice of the invention, including the use and performance of any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples include, but are not limited to, those having structural elements that do not differ from the literal language of the claims. or It is intended that the claims include equivalent structural elements that do not differ substantially from the literal language of the claims. [Explanation of symbols]

[0063] 10 Gas Turbine System 12 Compressor section, compressor part 14 Incoming Air Flow 16 Compressed Air 18 Combustor Section 20 fuel 22 Combustion gas 24 Turbine Section 26 rotor, shaft 28 External Load 30 steps 32 Turbine Blade 34 steps 36 Nozzles 38 Turbine casing 40 Airfoil 42 Outer Platform 44 Inner Platform 46 Airfoil 48 Tip part 50 Platform 52 Turbine shroud, stage 54 Turbine shroud 56 Extension 58 Opening 60 Main Unit 62 Support part 64 Middle part 66 Seal part 68 Front end 70 Front Hook 72 Rear end 74 Back Hook 76 First Surface 78 Second Surface 80 Slash Surface 82 Slash Surface 84 Non-straight segments 86 Anterior Segment 88 Hot Gas Path (HGP) Surfaces 90 Inlet opening 92 Cooling circuit, cooling system 94 Measuring plate 100 Plenum 102 first cooling passage 104 Front end 106 Rear end 108 Central part 110 Front 112 Posterior part 114 Ribs 116 Impingement opening 118 Second Cooling Passage 120 Ribs 122 Impingement opening 124 Front exhaust hole, exhaust channel 126 3rd Cooling Passage 128 Ribs 130 Impingement opening 132 Rear exhaust port 140 Cooling fluid 200 Flow control system, flow monitoring system 202 Cooling fluid 204 Pressure Operated Switch 206 Pneumatic Circuit 208 Air pressure passage 210 Fluid outlet 212 Fluid inlet 214 Piston 216 Disk 218 Seat area 220 Actuation element 304 Pressure Operated Switch 306 Transformable Chamber 308 Port 310 First Electrical Contact 312 Second Electrical Contact 314 First Surface 316 Controller 318 Extraction Port 900 Additive Manufacturing System, AM System 902 Object 904 Additive Manufacturing (AM) Control System 906 AM Printer 910 Chamber 912 Applicator 914 Raw materials 915 Build Plate 916 Laser or Electron Beam 918 Building Platform 920 Code 930 Computer 932 Memory 934 Processor 936 Input / Output (I / O) Interface 938 Bus 940 External I / O Devices / Resources 942 Storage Systems

Claims

1. 1. A flow regulation system (200) for increasing a flow rate of a cooling fluid (140, 202) supplied to a cooling system (92) of a component of a gas turbine system (10), the flow regulation system (200) comprising: a pneumatic circuit (206) embedded within a section of the component, the pneumatic circuit (206) including a set of interconnected pneumatic passages (208) and isolated from a cooling fluid (140, 202) flowing through the cooling system (92); a pressure activated switch (204, 304) fluidly coupled to the pneumatic circuit (206), the pressure activated switch (204, 304) actuating in response to the formation of a crack in the section of the component and exposure of at least one of the pneumatic passages (208) of the pneumatic circuit (206) embedded within the section of the component; Equipped with A flow regulation system (200), wherein the actuation of the pressure activated switch (204, 304) increases the flow rate of the cooling fluid (140, 202) supplied to the cooling system (92) of the component.

2. 2. The flow regulating system of claim 1, wherein the pressure activated switch is actuated in response to a drop in pressure in the pneumatic circuit due to the exposure of at least one of the pneumatic passages of the pneumatic circuit embedded within the section of the component.

3. 2. The flow regulating system of claim 1, wherein the cooling system of the component receives a flow of cooling fluid, and the actuation of the pressure activated switch causes a separate auxiliary flow of cooling fluid to enter the cooling system of the component.

4. 3. The flow regulating system of claim 2, wherein the pressure activated switch includes a fluid inlet fluidly coupled to a pressurized source of cooling fluid and a fluid outlet fluidly coupled to the cooling system of the component.

5. The flow regulation system (200) of claim 4, wherein the pressurized source of cooling fluid (140, 202) comprises a compressor discharge chamber (910) of a compressor component of the gas turbine system (10).

6. The flow regulating system (200) of claim 5, wherein pressure in the compressor discharge chamber (910) activates the pressure activated switch (204, 304) in response to a drop in pressure in the pneumatic circuit (206).

7. The flow regulation system (200) of claim 1, wherein the component of the gas turbine system (10) comprises a turbine shroud (54).

8. The flow regulation system (200) of claim 1, wherein the increase in flow rate of the cooling fluid (140, 202) provided to the cooling system (92) of the component prevents or inhibits oxidation adjacent the crevices of the component.

9. The pressure-activated switch (204, 304) a compressible chamber (910) exposed to pressure in a compressor discharge chamber (910) of a compressor component of the gas turbine system (10); a first electrical contact (310) on a first side of the compressible chamber (910); a second electrical contact (312) on a second opposing side of the compressible chamber (910); The flow regulation system (200) of claim 2, comprising:

10. 10. The flow regulating system of claim 9, further comprising a port extending through the first electrical contact, the port fluidly coupling an interior of the compressible chamber to the pneumatic circuit embedded within the section of the component.

11. 11. The flow regulating system of claim 10, wherein pressure in the compressor discharge chamber of the compressor component displaces the second electrical contact toward and into contact with the first electrical contact in response to a drop in pressure in the pneumatic circuit.

12. 12. The flow regulating system of claim 11, further comprising a controller for increasing a flow of cooling fluid through an extraction port of a compressor section of the gas turbine system into the cooling system of the component in response to the second electrical contact contacting the first electrical contact.

13. The flow regulation system (200) of any of the preceding claims, wherein the section of the component is a hot gas path surface (88) and the pneumatic circuit (206) is embedded in the hot gas path surface (88).

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