Patent Description:
Gas turbine systems are one example of turbomachines widely utilized in fields such as power generation. A conventional gas turbine system generally includes a compressor section, a combustor section, and a turbine section. During operation of a gas turbine system, various components in the system, such as turbine blades, nozzle airfoils, and shroud segments are subjected to high temperature gas flows, which can cause the components to fail. Since higher temperature flows generally result in increased performance, efficiency, and power output of a gas turbine system, it is advantageous to cool the components that are subjected to high temperature gas flows to allow the gas turbine system to operate at increased temperatures and to extend the lifetime of the components of a gas turbine system.

Cooling (e.g., convection cooling, impingement cooling, etc.) is often provided by directing a flow of a cooling fluid through internal passages formed in the components of the gas turbine system. In many cases, the cooling fluid is provided by bleeding off a portion of the air discharged by the compressor section of the gas turbine system.

A thermal barrier coating (TBC) is often applied to the components of a gas turbine system to provide a protective heat shield, prevent damage due to high temperatures, and extend component life by reducing oxidation and thermal fatigue. Spallation of the TBC is a common issue in gas turbine systems. When the TBC spalls, portions of the TBC may crack and break off of a component, exposing underlying surfaces to high temperatures and damage (e.g., due to oxidation).

<CIT> discloses an apparatus for controlling oxidation in a vicinity of a breach in a component of a gas turbine system. If a damage occurs through a portion of an annular runner, a pressure sensor monitors a change in the pressure in a buffer chamber. A controller is coupled to a valve to open and close the valve in response to signals received from the pressure sensor to modulate the pressurized cooling air directed into the buffer chamber and to provide supplemental flow or an increased flow of cooling fluid impeding or arresting oxidation in the vicinity of the breach in the component.

<CIT> describes a detector device for monitoring the operational characteristics of a compressor or turbine rotor, including a detector head communicating with a pressure-actuated switch via a conduit. The pressure-actuated switch senses the difference in pressure between the high pressure air surrounding the detector head and the ambient pressure surrounding the pressure-actuated switch. The detector member comprises a hollow body member provided with a frangible probe which is positioned in close proximity to a rim of a turbine disc. In operation, if the eccentricity occurring in the turbine disc exceeds a predetermined value, the rim of the turbine disc will come into contact with the frangible probe and fracture it, allowing the high pressure cooling air passing around the turbine disc to enter the hollow body member and in turn operate the pressure-responsive switch via conduit. The pressure-actuated switch provides a signal to an indicating device via a wire, and the indicating device may operate a warning device.

The invention provides a flow regulating system for increasing a flow of cooling fluid supplied to a cooling system of a component of a gas turbine system having the features of independent claims <NUM> and <NUM>. Especially preferred embodiments of the invention are defined in the dependent claims.

The illustrative aspects of the present disclosure solve the problems herein described and/or other problems not discussed.

These and other features of this disclosure will be more readily understood from the following detailed description of the various aspects of the disclosure taken in conjunction with the accompanying drawings that depict various embodiments of the disclosure.

Reference will now be made in detail to representative embodiments illustrated in the accompanying drawings. It should be understood that the following descriptions are not intended to limit the embodiments to one preferred embodiment. To the contrary, it is intended to cover alternatives, modifications, and equivalents as can be included within the scope of the described embodiments as defined by the appended claims.

As an initial matter, in order to clearly describe the current disclosure, it will become necessary to select certain terminology when referring to and describing relevant machine components within the scope of this disclosure. When doing this, if possible, common industry terminology will be used and employed in a manner consistent with its accepted meaning. Unless otherwise stated, such terminology 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 often a particular component may be referred to using several different or overlapping terms. What may be described herein as being a single part may include and be referenced in another context as consisting of multiple components. Alternatively, what may be described herein as including multiple components may be referred to elsewhere as a single part.

In addition, several descriptive terms may be used regularly herein, and it should prove helpful to define these terms at the onset of this section. These terms and their definitions, unless stated otherwise, are as follows. As used herein, "downstream" and "upstream" are terms that indicate a direction relative to the flow of a fluid, such as the working fluid through the turbine or, for example, the flow of air through the combustor or coolant through one of the turbine's component systems. The term "downstream" corresponds to the direction of flow of the fluid, and the term "upstream" refers to the direction opposite to the flow. The terms "forward" and "aft," without any further specificity, refer to directions, with "forward" referring to the front or compressor end of the engine, and "aft" referring to the rearward or turbine end of the engine. Additionally, the terms "leading" and "trailing" may be used and/or understood as being similar in description as the terms "forward" and "aft," respectively. It is often required to describe parts that are at differing radial, axial and/or circumferential positions. The "A" axis represents an axial orientation. As used herein, the terms "axial" and/or "axially" refer to the relative position/direction of objects along axis A, which is substantially parallel with the axis of rotation of the gas turbine system (in particular, the rotor section). As further used herein, the terms "radial" and/or "radially" refer to the relative position/direction of objects along a direction "R" (see, <FIG>), which is substantially perpendicular with axis A and intersects axis A at only one location. Finally, the term "circumferential" refers to movement or position around axis A (e.g., direction "C").

In various embodiments, components described as being "fluidly coupled" to or "in fluid communication" with one another can be joined along one or more interfaces. In some embodiments, these interfaces can include junctions between distinct components, and in other cases, these interfaces can include a solidly and/or integrally formed interconnection. That is, in some cases, components that are "coupled" to one another can be simultaneously formed to define a single continuous member. However, in other embodiments, these coupled components can be formed as separate members and be subsequently joined through known processes (e.g., fastening, ultrasonic welding, bonding).

When an element or layer is referred to as being "on", "engaged to", "connected to" or "coupled to" another element, it may be directly on, engaged, connected or coupled to the other element, or intervening elements may be present. In contrast, when an element is referred to as being "directly on," "directly engaged to", "directly connected to" or "directly coupled to" another element, there may be no intervening elements or layers present.

<FIG> depicts a schematic diagram of a gas turbine system <NUM> according to various embodiments. As shown, the gas turbine system <NUM> includes a compressor section <NUM> for compressing an incoming flow of air <NUM> and for delivering a flow of compressed air <NUM> to a combustor section <NUM>. The combustor section <NUM> mixes the flow of compressed air <NUM> with a pressurized supply of fuel <NUM> and ignites the mixture to create a flow of combustion gases <NUM>. Although only a single combustor section <NUM> is shown, the gas turbine system <NUM> may include any number of combustor sections <NUM>. The flow of combustion gases <NUM> is in turn delivered to a turbine section <NUM>. The flow of combustion gases <NUM> drives the turbine section <NUM> to produce mechanical work. The mechanical work produced in the turbine section <NUM> may drive the compressor section <NUM> via a shaft <NUM> and may be used to drive an external load <NUM>, such as an electrical generator and/or the like.

<FIG> depicts a side view of a portion of a turbine section <NUM> of a gas turbine system, including at least one stage <NUM> of turbine blades <NUM> (one shown) and at least one stage <NUM> of nozzles <NUM> (one shown) positioned within a casing <NUM> of the turbine section <NUM> ("turbine casing <NUM>"). Each stage <NUM> of turbine blades <NUM> includes a plurality of turbine blades <NUM> that are coupled to and positioned circumferentially about the rotor <NUM>, and which are driven by the combustion gases <NUM>. Each stage <NUM> of nozzles <NUM> includes a plurality of nozzles <NUM> that are coupled to and positioned circumferentially about the turbine casing <NUM> of the turbine section <NUM>.

In the embodiment shown in <FIG>, each nozzle <NUM> includes an airfoil <NUM> positioned between an outer platform <NUM> and an inner platform <NUM>. Similar to the nozzles <NUM>, each turbine blade <NUM> of the turbine section <NUM> includes an airfoil <NUM> extending radially from the rotor <NUM>. Each airfoil <NUM> includes a tip portion <NUM> and a platform <NUM> positioned opposite the tip portion <NUM>.

The turbine blades <NUM> and the nozzles <NUM> may be positioned axially adjacent to one another within the turbine casing <NUM>. In <FIG>, for example, the nozzles <NUM> are shown positioned axially adjacent and downstream of the turbine blades <NUM>. The turbine section <NUM> may include a plurality of stages <NUM> of turbine blades <NUM> and a plurality of stages <NUM> of nozzles <NUM>, positioned axially throughout the turbine casing <NUM>.

The turbine section <NUM> of the gas turbine system <NUM> may include a plurality of stages <NUM> of shrouds <NUM> (one stage shown in <FIG>) positioned axially throughout the turbine casing <NUM>. In <FIG>, for example, the stage <NUM> of shrouds <NUM> is shown positioned radially adjacent to and substantially surrounding or encircling the stage <NUM> of turbine blades <NUM>. The stage <NUM> of shrouds <NUM> may also be positioned axially adjacent and/or upstream of the stage <NUM> of nozzles <NUM>. Further, the stage <NUM> of shrouds <NUM> may be positioned between two adjacent stages <NUM> of nozzles <NUM> located on opposing sides of a stage <NUM> of turbine blades <NUM>. The stage <NUM> of shrouds <NUM> may be coupled about the turbine casing <NUM> using a set of extensions <NUM>, each including an opening <NUM> configured to receive a corresponding section of a shroud <NUM>.

An isometric view of a turbine shroud <NUM> is depicted in <FIG> and a cross-sectional view of the turbine shroud <NUM> is depicted in <FIG>. As shown, the turbine shroud <NUM> includes a body <NUM>. The body <NUM> of the turbine shroud <NUM>, and various other components and/or features of the turbine shroud <NUM>, may be formed using any suitable technique, including an additive manufacturing process. For example, the turbine shroud <NUM> including body <NUM> may be formed by direct metal laser melting (DMLM) (also referred to as selective laser melting (SLM)), direct metal laser sintering (DMLS), electronic beam melting (EBM), stereolithography (SLA), binder jetting, or any other suitable additive manufacturing process.

The body <NUM> of the turbine shroud <NUM> includes a support portion <NUM>, an intermediate portion <NUM>, and a seal portion <NUM>. The support portion <NUM> is coupled directly to and/or aids in the coupling of the turbine shroud <NUM> to the turbine casing <NUM> and/or extension <NUM> (see, <FIG>). The support portion <NUM> includes a forward end <NUM> including at least one forward hook <NUM>, an aft end <NUM> including at least one aft hook <NUM>, a first surface <NUM>, and a second surface <NUM>. The intermediate portion <NUM> includes various features of the body <NUM> between opposing slash faces <NUM>, <NUM> of the body <NUM>, including a non-linear segment <NUM> and a forward segment <NUM>. The forward segment <NUM> may be used, for example, to form a seal within the turbine section <NUM>, define a hot gas flow path of combustion gases <NUM> flowing through the turbine section <NUM>, and/or secure nozzles <NUM> within the turbine casing <NUM>. The seal portion <NUM> may at least partially define the flow path of combustion gases <NUM> flowing through turbine section <NUM>. The seal portion <NUM> includes a hot gas path (HGP) surface <NUM> that may be positioned adjacent the hot gas flow path of combustion gases <NUM> within the turbine section <NUM>.

The body <NUM> of the turbine shroud <NUM> further includes at least one inlet opening <NUM>, formed in and/or through the first surface <NUM> of the support portion <NUM>, between the forward end <NUM> and the aft end <NUM> of the body <NUM>. The inlet opening <NUM> is in fluid communication with a cooling circuit <NUM> (<FIG>) formed through and/or included within the support portion <NUM>, intermediate portion <NUM>, and seal portion <NUM> of the body <NUM>.

The turbine shroud <NUM> may also include a set of metering plates <NUM> (shown in phantom in <FIG>) coupled to the first surface <NUM> of the support portion <NUM> of the body <NUM>. Each metering plate <NUM> (only one shown) may be affixed to the first surface <NUM>, over and/or at least partially covering a respective inlet opening <NUM>. The metering plate <NUM> allows a predetermined flow of cooling fluid to enter the cooling circuit <NUM> via the inlet opening <NUM>.

Various plenum(s) and/or cooling passage(s) of the turbine shroud <NUM> are depicted in <FIG>, which is a cross-sectional view of the turbine shroud <NUM> of <FIG>. As shown, the turbine shroud <NUM> includes at least one plenum <NUM>, which may be formed and/or extend through a portion of the body <NUM> of the turbine shroud <NUM>. More specifically, the plenum <NUM> may extend (radially) through at least a portion of the support portion <NUM>, intermediate portion <NUM>, and the seal portion <NUM> of the body <NUM> of the turbine shroud <NUM>. One or more portions of the plenum <NUM> formed within the intermediate portion <NUM> and the seal portion <NUM> of the body <NUM> may extend between and/or adjacent the opposing slash faces <NUM>, <NUM>. Although only a single plenum <NUM> is shown, it is understood that the turbine shroud <NUM> may include additional plenums.

The plenum <NUM> is fluidly coupled to and/or in direct fluid communication with the inlet opening(s) <NUM> formed in the support portion <NUM> of the body <NUM>. As discussed herein, the plenum <NUM> is configured to receive a supply of cooling fluid <NUM> (e.g., compressor discharge air) via the inlet opening(s) <NUM>, and may provide the cooling fluid <NUM> to distinct cooling passages formed in the turbine shroud <NUM> to cool the turbine shroud <NUM> during operation of the gas turbine system <NUM>.

As shown in <FIG>, the turbine shroud <NUM> includes a first cooling passage <NUM> formed, positioned, and/or extending within the body <NUM> of the turbine shroud <NUM> and in fluid communication with the plenum <NUM>. More specifically, the first cooling passage <NUM> may be positioned within and/or extend through the seal portion <NUM> of the body <NUM> of the turbine shroud <NUM>, between and/or adjacent a forward end <NUM> and an aft end <NUM> of the body <NUM>. Additionally, the first cooling passage <NUM> may extend through the seal portion <NUM> of the body <NUM> between and/or adjacent the opposing slash faces <NUM>, <NUM>. The first cooling passage <NUM> may also be positioned within the seal portion <NUM> radially between the plenum <NUM> and the HGP surface <NUM> of the seal portion <NUM>.

The first cooling passage <NUM> may include a plurality of distinct segments, sections, and/or parts. For example, the first cooling passage <NUM> is shown as including a central part <NUM> positioned and/or extending between a forward part <NUM>, and an aft part <NUM>. As shown in <FIG>, the central part <NUM> of the first cooling passage <NUM> is centrally formed and/or positioned between the forward end <NUM> and the aft end <NUM> of the seal portion <NUM>. The forward part <NUM> of the first cooling passage <NUM> is formed and/or positioned directly adjacent the forward end <NUM> of the seal portion <NUM>, and axially adjacent and/or axially upstream of the central part <NUM>. Similarly, the aft part <NUM> of the first cooling passage <NUM> is formed and/or positioned directly adjacent the aft end <NUM> of the seal portion <NUM>, opposite the forward part <NUM>. The central part <NUM> may be radially aligned with the axial portion of the HGP surface <NUM> of the seal portion <NUM> that requires the most cooling and/or demands the largest heat exchange within the turbine shroud <NUM> to improve the operational efficiency of the turbine section <NUM> and/or the operational life of the turbine shroud <NUM> within the turbine section <NUM>.

The plenum <NUM> and the first cooling passage <NUM> are separated by a rib <NUM>. The rib <NUM> may extend within the body <NUM> of the turbine shroud <NUM> between the opposing slash faces <NUM>, <NUM>. A plurality of impingement openings <NUM> are provided through the rib <NUM>. The impingement openings <NUM> fluidly couple the plenum <NUM> and the central part <NUM> of the first cooling passage <NUM>. During operation of the gas turbine system <NUM>, cooling fluid <NUM> flows from the plenum <NUM> through the plurality of impingement openings <NUM> into the central portion <NUM> of the first cooling passage <NUM>.

In addition to first cooling passage <NUM>, the body <NUM> of the turbine shroud <NUM> includes a second cooling passage <NUM>. As depicted in <FIG>, the second cooling passage <NUM> extends within the body <NUM> of the turbine shroud <NUM> adjacent the forward end <NUM> of the seal portion <NUM>, and may extend within the seal portion <NUM> between the opposing slash faces <NUM>, <NUM>. The second cooling passage <NUM> is positioned adjacent to and upstream of the central part <NUM> of the first cooling passage <NUM>, and is positioned radially inward from the forward part <NUM> of the first cooling passage <NUM>. Further, the second cooling passage <NUM> is formed or positioned between the forward part <NUM> of the first cooling passage <NUM> and the HGP surface <NUM> of the seal portion <NUM>.

A rib <NUM> separates the second cooling passage <NUM> from the forward part <NUM> of the first cooling passage <NUM>. The rib <NUM> may extend within the body <NUM> of the turbine shroud <NUM> between the opposing slash faces <NUM>, <NUM>. The second cooling passage <NUM> is in direct fluid communication with the forward part <NUM> of the first cooling passage <NUM>. A plurality of impingement openings <NUM>, formed through the rib <NUM>, fluidly couple the forward part <NUM> of the first cooling passage <NUM> and the second cooling passage <NUM>. During operation of the gas turbine system <NUM> (see, <FIG>), cooling fluid <NUM> flowing through the forward part <NUM> of the first cooling passage <NUM> passes or flows from the forward part <NUM> of the first cooling passage <NUM> through the impingement openings <NUM> to second cooling passage <NUM>.

The body <NUM> of the turbine shroud <NUM> further includes a plurality of forward exhaust holes <NUM> in fluid communication with the second cooling passage <NUM>. During operation, the plurality of forward exhaust holes <NUM> discharge cooling fluid <NUM> from the second cooling passage <NUM> into the hot gas flow path of combustion gases <NUM> flowing through the turbine section <NUM> of the gas turbine system <NUM>.

A third cooling passage <NUM> is provided in the seal portion <NUM> of the body <NUM> of the turbine shroud <NUM>. As shown in <FIG>, the third cooling passage <NUM> is located adjacent the aft end <NUM> of the seal portion <NUM>. As with the first and second cooling passages <NUM>, <NUM>, the third cooling passage <NUM> may also extend within the seal portion <NUM> between the opposing slash faces <NUM>, <NUM>. The third cooling passage <NUM> is positioned adjacent to and downstream of the central part <NUM> of the first cooling passage <NUM>, and is positioned radially inward from the aft part <NUM> of the first cooling passage <NUM>, adjacent the HGP surface <NUM> of the seal portion <NUM>.

A rib <NUM> separates the third cooling passage <NUM> from the aft part <NUM> of the first cooling passage <NUM>. The rib <NUM> may extend within the body <NUM> of the turbine shroud <NUM> between the opposing slash faces <NUM>, <NUM>. A plurality of impingement openings <NUM>, formed through the rib <NUM>, fluidly couple the aft part <NUM> of the first cooling passage <NUM> and the third cooling passage <NUM>. During operation of the gas turbine system <NUM> (see, <FIG>), cooling fluid <NUM> flowing through the aft part <NUM> of the first cooling passage <NUM> passes or flows from the aft part <NUM> of the first cooling passage <NUM> through the impingement openings <NUM> to the third cooling passage <NUM>.

The body <NUM> of the turbine shroud <NUM> further includes a plurality of aft exhaust holes <NUM> in fluid communication with the third cooling passage <NUM>. The plurality of aft exhaust holes <NUM> are configured to discharge cooling fluid <NUM> from the third cooling passage <NUM> into the hot gas flow path of combustion gases <NUM> flowing through the turbine section <NUM> of the gas turbine system <NUM>.

During operation of gas turbine system <NUM>, cooling fluid <NUM> (e.g., compressor discharge air) flows under pressure through the body <NUM> to cool the turbine shroud <NUM>. More specifically, as the turbine shroud <NUM> is exposed to the combustion gases <NUM> flowing through the hot gas flow path of the turbine section <NUM> during operation of gas turbine system <NUM>, cooling fluid <NUM> is provided to various cooling features (e.g., plenum <NUM>, cooling passages <NUM>, <NUM>, <NUM>, exhaust channels <NUM>, <NUM>, etc.) within the body <NUM> to cool the turbine shroud <NUM>. In a non-limiting example, the cooling fluid <NUM> flows into the plenum <NUM> in the body <NUM> of the turbine shroud <NUM> through the inlet opening(s) <NUM> formed in first surface <NUM> of the support portion <NUM> of the body <NUM> of the turbine shroud <NUM>. Additionally where the turbine shroud <NUM> includes metering plate(s) <NUM> affixed to first surface <NUM>, over and/or at least partially covering the inlet opening(s) <NUM>, the metering plate(s) <NUM> regulate the amount and pressure of a fixed flow of the cooling fluid <NUM> flowing through the inlet opening(s) <NUM> into the plenum(s) <NUM>.

The cooling fluid <NUM> is configured to flow from the inlet opening(s) <NUM>, through the plenum <NUM>, and radially toward the cooling passages <NUM>, <NUM>, <NUM> formed within the seal portion <NUM> of the body <NUM> of the turbine shroud <NUM>. More specifically, the cooling fluid <NUM> provided to the plenum <NUM> flows radially through the plenum <NUM> toward the rib <NUM>, and subsequently through the impingement openings <NUM> in the rib <NUM> to the central part <NUM> of the first cooling passage <NUM>. From the central part <NUM> of the first cooling passage <NUM>, the cooling fluid <NUM> flows axially into the forward part <NUM> and aft part <NUM> of the first cooling passage <NUM>.

The portion of the cooling fluid <NUM> flowing into the forward part <NUM> of the first cooling passage <NUM> flows through the impingement openings <NUM> formed in the rib <NUM> into the second cooling passage <NUM>. Similarly, the portion of the cooling fluid <NUM> flowing into the aft part <NUM> of the first cooling passage <NUM> flows through the impingement openings <NUM> formed in the rib <NUM> into the third cooling passage <NUM>.

From the second cooling passage <NUM>, a portion of the cooling fluid <NUM> may flow out of the body <NUM> of the turbine shroud <NUM> through the exhaust holes <NUM>. Additionally, a portion of the cooling fluid <NUM> in the third cooling passage <NUM> may flow out of the body <NUM> through the exhaust holes <NUM>. The remaining cooling fluid <NUM> not exhausted via the exhaust holes <NUM>, <NUM> may be provided to other cooling features in the body <NUM> of the turbine shroud <NUM>.

A breach may form in the HGP surface <NUM> of the turbine shroud <NUM> as a result of oxidation following TBC spallation. The breach may increase in size due to continued oxidation of the HGP surface <NUM> in the vicinity of the breach, reducing the operational lifetime of the turbine shroud <NUM>.

According to embodiments, a flow regulating system <NUM> is provided for delivering a supplemental flow of cooling fluid <NUM> to the cooling circuit <NUM> of the turbine shroud <NUM> in response to the formation of a breach in the HGP surface <NUM>. In particular, in response to the formation of a breach in the HGP surface <NUM>, the flow regulating system <NUM> is configured to direct the supplemental flow of cooling fluid <NUM> into the cooling circuit <NUM> (e.g., into the plenum <NUM>) in the body <NUM> of the turbine shroud <NUM>. The flow regulating system <NUM> is used, for example as depicted in <FIG>, to provide a supplemental flow of cooling fluid <NUM> to the cooling circuit <NUM> of a turbine shroud <NUM> that has its cooling fluid <NUM> fed directly from the discharge of the compressor section <NUM> (<FIG>) of the gas turbine system <NUM> (e.g., <NUM>st stage or above). According to other embodiments, the flow regulating system <NUM> is used, for example as depicted in <FIG> and <FIG>, to increase the flow of cooling fluid <NUM> to the cooling circuit <NUM> of a turbine shroud <NUM> that receives its cooling fluid <NUM> through an extraction port of the compressor section <NUM> of the gas turbine system <NUM> (e.g., <NUM>nd stage or above). Although described herein in conjunction with a turbine shroud <NUM>, it should be noted that the flow regulating system <NUM> may be configured for use with other components of a gas turbine system <NUM> that may be subject to oxidation. For example, the flow regulating system <NUM> may be configured to provide a supplemental supply of cooling fluid (or an increased flow of cooling fluid) to the cooling system of a component such as a turbine blade, nozzle, and/or the like, in response to the formation of a breach in a HGP surface of the component.

According to various embodiments, the flow regulating system <NUM> includes a pressure-actuated switch <NUM> (<FIG>) or <NUM> (<FIG> and <FIG>) and a pneumatic circuit <NUM> (<FIG> and <FIG>)including a set of interconnected pneumatic passages <NUM> entirely embedded within the HGP surface <NUM>. At least one of the pneumatic passages <NUM> is fluidly coupled to the pressure-actuated switch <NUM>, <NUM>. As shown in <FIG> and <FIG>, the pneumatic passages <NUM> may be provided in a grid-like pattern within the HGP surface <NUM>. Many other configurations may also be used. For example, the pneumatic passages <NUM> may be provided in a sinusoidal configuration, a cross-hatched configuration, a spiral configuration, a rectangular grid, etc. In general, the arrangement and spacing of the pneumatic passages <NUM> within the HGP surface <NUM> is such that even a relatively small breach in the HGP surface <NUM> will result in the exposure of one or more of the pneumatic passages <NUM> in the HGP surface <NUM>. In the absence of a breach, none of the pneumatic passages <NUM> of the pneumatic circuit <NUM> are exposed.

During normal operation (e.g., the absence of breach in the HGP surface <NUM>), the pressure within the pneumatic passages <NUM> of the pneumatic circuit <NUM> is sufficient to maintain the pressure-actuated switch <NUM>, <NUM> in a non-actuated state. When a breach forms in the HGP surface <NUM>, resulting in the exposure of one or more of the pneumatic passages <NUM> of the pneumatic circuit <NUM> embedded within the HGP surface <NUM>, the pressure within the pneumatic passages <NUM> of the pneumatic circuit <NUM> drops, resulting in an actuation of the pressure-actuated switch <NUM>, <NUM>. When actuated, the pressure-actuated switch <NUM>, <NUM> allows a supplemental supply of cooling fluid <NUM> (or an increased flow of cooling fluid <NUM>) to flow into the cooling system <NUM> of the turbine shroud <NUM>. Advantageously, the increased flow of cooling fluid into the cooling system <NUM> is such that the oxidation of the HGP surface <NUM> in the vicinity of the breach is impeded or arrested.

A pressure-actuated switch <NUM> according to embodiments is depicted in <FIG>. The pressure-actuated switch <NUM> is configured to selectively provide a supplemental supply of cooling fluid <NUM> to the cooling system <NUM> of the turbine shroud <NUM> in response to the formation of a breach in the HGP surface <NUM>. The pressure-actuated switch <NUM> includes a fluid outlet <NUM> and a fluid inlet <NUM> that is fluidly coupled to a pressurized source of cooling fluid. In <FIG>, for example, the fluid inlet <NUM> of the pressure-actuated switch <NUM> is fluidly coupled to a compressor discharge chamber (CDC), which receives a flow of compressed air <NUM> (<FIG>) from the compressor portion <NUM> of the gas turbine system <NUM>.

The pressure-actuated switch <NUM> further includes a piston <NUM> and a disc <NUM> coupled to the piston <NUM> that is configured to mate with a seat <NUM>. A biasing element <NUM>, such as a spring or the like, biases the disc <NUM> toward the seat <NUM>. A distal end of the piston <NUM> is fluidly coupled to the pneumatic circuit <NUM> via one or more of the pneumatic passages <NUM>.

In the absence of a breach in the HGP surface <NUM>, as shown in <FIG>, the pressure within the pneumatic passages <NUM> together with the biasing force applied by the biasing element <NUM> is greater than the pressure within the compressor discharge chamber. As such, the disc <NUM> coupled to the piston <NUM> is forced against the seat <NUM>, closing the pressure-actuated switch <NUM>. This prevents the supplemental supply of cooling fluid <NUM> from flowing from the compressor discharge chamber into the cooling system <NUM> of the turbine shroud <NUM>.

The pressure-actuated switch <NUM> is actuated in response to the formation of a breach (e.g., due to oxidation) in the HGP surface <NUM> and an exposure of at least a portion of the pneumatic passages <NUM> of the pneumatic circuit <NUM> embedded within the HGP surface <NUM>. The exposure causes a loss of pressure within the pneumatic passages <NUM> of the pneumatic circuit <NUM>. As a result, as shown in <FIG>, the pressure within the compressor discharge chamber is now greater than the biasing force applied by the biasing element <NUM>, forcing the disc <NUM> coupled to the piston <NUM> away from the seat <NUM> and fluidly coupling the fluid inlet <NUM> to the fluid outlet <NUM> of the pressure-actuated switch <NUM>. A supplemental supply of cooling fluid <NUM> can now flow from the compressor discharge chamber through the pressure-actuated switch <NUM> into the cooling system <NUM> of the turbine shroud <NUM>. The increased flow of cooling fluid into the cooling system <NUM> impedes or arrests the oxidation of the HGP surface <NUM> in the vicinity of the breach.

A pressure-actuated switch <NUM> according to other embodiments is depicted in <FIG> and <FIG>. The pressure-actuated switch <NUM> is configured to selectively trigger an increase in the flow of cooling fluid <NUM> provided to the cooling system <NUM> of the turbine shroud <NUM> in response to the formation of a breach in the HGP surface <NUM>. The pressure-actuated switch <NUM> may be used, for example, in a turbine shroud <NUM> that receives its cooling fluid <NUM> via an extraction port of the compressor section <NUM> of the gas turbine system <NUM> (e.g., <NUM>nd stage or above).

The pressure-actuated switch <NUM> may be positioned on the body <NUM> of the turbine shroud <NUM>. The pressure-actuated switch <NUM> includes a deformable chamber <NUM> that is fluidly coupled via a port <NUM> to one or more of the pneumatic passages <NUM> of the pneumatic circuit <NUM>. The port <NUM> passes through a first electrical contact <NUM> positioned on a first surface of the deformable chamber <NUM>. A second electrical contact <NUM> is provided on a second, opposing surface <NUM> of the deformable chamber <NUM>. The second surface <NUM> of the deformable chamber <NUM> is exposed to the pressure within the compressor discharge chamber, which receives a flow of compressed air <NUM> (<FIG>) from the compressor portion <NUM> of the gas turbine system <NUM>.

In the absence of a breach in the HGP surface <NUM>, as shown in <FIG>, the pressure within the pneumatic passages <NUM> (and within the deformable chamber <NUM>) is higher than the pressure within the compressor discharge chamber. The higher pressure in the deformable chamber <NUM> causes the deformable chamber <NUM> to expand, which prevents the second electrical contact <NUM> from coming into contact with the first electrical contact <NUM>. The pressure-actuated switch <NUM> is thus in an non-actuated state. The first and second electrical contacts <NUM>, <NUM> are electrically connected to a controller <NUM>.

The pressure-actuated switch <NUM> is activated in response to the formation of a breach (e.g., due to oxidation) in the HGP surface <NUM> and an exposure of at least a portion of the pneumatic passages <NUM> of the pneumatic circuit <NUM> embedded within the HGP surface <NUM>. The exposure causes a loss of pressure within the pneumatic passages <NUM> of the pneumatic circuit <NUM>. As a result, the compressor discharge chamber pressure is now greater than the pressure within the deformable chamber <NUM>. As shown in <FIG>, this results in the collapse of the deformable chamber <NUM>. As a consequence, the second electrical contact <NUM> is displaced toward and against the first electrical contact <NUM>, completing an electrical circuit and activating the pressure-actuated switch <NUM>.

In response to the activation of the pressure-actuated switch <NUM>, the controller <NUM> increases the amount of cooling fluid <NUM> flowing through the extraction port <NUM> into the cooling system <NUM> of the turbine shroud <NUM>. The increased flow of cooling fluid <NUM> may be provided, for example, by enlarging the size of the opening of the extraction port <NUM> as shown in phantom in <FIG> in response to a signal from the controller <NUM>. The increased flow of cooling fluid <NUM> into the cooling system <NUM> impedes or arrests the oxidation of the HGP surface <NUM> in the vicinity of the breach.

A flow diagram of a method for impeding and/or arresting oxidation, which as such is not claimed, is depicted in <FIG>, and is described with reference to <FIG>. At S1, the pressure within the pneumatic passages <NUM> of the pneumatic circuit <NUM> is monitored by a flow monitoring system <NUM>. In response to a drop in pressure in the pneumatic circuit <NUM> (YES at S2), flow passes to S3. In the absence of a drop in pressure in the pneumatic circuit <NUM> (NO at S2), flow passes back to S1.

At S3, a pressure-actuated switch <NUM>, <NUM> is actuated in response to the drop in pressure in the pneumatic circuit <NUM>. At S4, a supplemental flow of cooling fluid <NUM> or an increased flow of cooling fluid <NUM> is provided to the cooling system <NUM> of the turbine shroud <NUM>. At S5, oxidation in the vicinity of the breach in the HGP surface <NUM> of the turbine shroud <NUM> is impeded or arrested due to the supplemental flow of cooling fluid <NUM> or the increased flow of cooling fluid <NUM>.

As detailed above, the body <NUM> of the turbine shroud <NUM>, and various other components and/or features of the turbine shroud <NUM>, including various components/portions of the flow regulating system <NUM> disclosed herein, may be formed using any suitable technique, including an additive manufacturing process. The additive manufacturing process may use any suitable material capable of withstanding the operational characteristics (e.g., exposure temperature, exposure pressure, and the like) experienced by the turbine shroud <NUM> within gas turbine system <NUM> during operation.

As used herein, additive manufacturing (AM) may include any process of producing an object through the successive layering of material rather than the removal of material, which is the case with conventional processes. Additive manufacturing can create complex geometries without the use of any sort of tools, molds or fixtures, and with little or no waste material. Instead of machining components from solid billets of plastic or metal, much of which is cut away and discarded, the only material used in additive manufacturing is what is required to shape the part. Additive manufacturing processes may 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 setting, DMLM or SLM have been found advantageous.

To illustrate an example of an additive manufacturing process, <FIG> shows a schematic/block view of an illustrative computerized additive manufacturing system <NUM> for generating an object <NUM>. In this example, the system <NUM> is arranged for DMLM. It is understood that the general teachings of the disclosure are equally applicable to other forms of additive manufacturing. The object <NUM> is illustrated as a turbine shroud <NUM> (see, <FIG>). The AM system <NUM> generally includes a computerized additive manufacturing (AM) control system <NUM> and an AM printer <NUM>. The AM system <NUM>, as will be described, executes code <NUM> that includes a set of computer-executable instructions defining the turbine shroud <NUM> to physically generate the object <NUM> using the AM printer <NUM>. Each AM process may use different raw materials in the form of, for example, fine-grain powder, liquid (e.g., polymers), sheet, etc., a stock of which may be held in a chamber <NUM> of the AM printer <NUM>. In the instant case, the turbine shroud <NUM> may be made of a metal or metal compound capable of withstanding the environment of a gas turbine system <NUM> (see, <FIG>). As illustrated, an applicator <NUM> may create a thin layer of raw material <NUM> spread out as the blank canvas on a build plate <NUM> of AM printer <NUM> from which each successive slice of the final object will be created. In other cases, the applicator <NUM> may directly apply or print the next layer onto a previous layer as defined by code <NUM>, e.g., where a metal binder jetting process is used. In the example shown, a laser or electron beam <NUM> fuses particles for each slice, as defined by code <NUM>, but this may not be necessary where a quick setting liquid plastic/polymer is employed. Various parts of the AM printer <NUM> may move to accommodate the addition of each new layer, e.g., a build platform <NUM> may lower and/or chamber <NUM> and/or applicator <NUM> may rise after each layer.

The AM control system <NUM> is shown implemented on a computer <NUM> as computer program code. To this extent, the computer <NUM> is shown including a memory <NUM>, a processor <NUM>, an input/output (I/O) interface <NUM>, and a bus <NUM>. Further, the computer <NUM> is shown in communication with an external I/O device/resource <NUM> and a storage system <NUM>. In general, the processor <NUM> executes computer program code, such as the AM control system <NUM>, that is stored in memory <NUM> and/or storage system <NUM> under instructions from code <NUM> representative of turbine shroud <NUM>, described herein. While executing computer program code, the processor <NUM> can read and/or write data to/from memory <NUM>, storage system <NUM>, I/O device <NUM>, and/or AM printer <NUM>. The bus <NUM> provides a communication link between each of the components in the computer <NUM>, and the I/O device <NUM> can comprise any device that enables a user to interact with computer <NUM> (e.g., keyboard, pointing device, display, etc.). The computer <NUM> is only representative of various possible combinations of hardware and software. For example, the processor <NUM> may comprise a single processing unit, or be distributed across one or more processing units in one or more locations, e.g., on a client and server. Similarly, the memory <NUM> and/or storage system <NUM> may reside at one or more physical locations. The memory <NUM> and/or storage system <NUM> can comprise any combination of various types of non-transitory computer readable storage medium including magnetic media, optical media, random access memory (RAM), read only memory (ROM), etc. The computer <NUM> can comprise any type of computing device such as a network server, a desktop computer, a laptop, a handheld device, a mobile phone, a pager, a personal data assistant, etc..

Additive manufacturing processes begin with a non-transitory computer readable storage medium (e.g., memory <NUM>, storage system <NUM>, etc.) storing code <NUM> representative of the turbine shroud <NUM>. For example, the code <NUM> may include a precisely defined 3D model of the turbine shroud <NUM> and can be generated from any of a large variety of well-known computer aided design (CAD) software systems such as AutoCAD®, TurboCAD®, DesignCAD 3D Max, etc. In this regard, the code <NUM> can take any now known or later developed file format. For example, the code <NUM> may be in the Standard Tessellation Language (STL) which was created for stereolithography CAD programs of 3D Systems, or an additive manufacturing file (AMF), which is an American Society of Mechanical Engineers (ASME) standard that is an extensible markup-language (XML) based format designed to allow any CAD software to describe the shape and composition of any three-dimensional object to be fabricated on any AM printer. The code <NUM> may be translated between different formats, converted into a set of data signals and transmitted, received as a set of data signals and converted to code, stored, etc., as necessary. The code <NUM> may be an input to system <NUM> and may come from a part designer, an intellectual property (IP) provider, a design company, the operator or owner of system <NUM>, or from other sources. In any event, the AM control system <NUM> executes the code <NUM>, dividing the turbine shroud <NUM> into a series of thin slices that it assembles using the AM printer <NUM> in successive layers of liquid, powder, sheet or other material. In the DMLM example, each layer is melted to the exact geometry defined by the code <NUM> and fused to the preceding layer. Subsequently, the turbine shroud <NUM> may be exposed to any variety of finishing processes, e.g., those described herein for re-contouring or other minor machining, sealing, polishing, etc..

Claim 1:
A flow regulating system (<NUM>) for increasing a flow of cooling fluid (<NUM>, <NUM>) supplied to a cooling system of a component of a gas turbine system (<NUM>), comprising:
a pneumatic circuit (<NUM>) embedded within a section of the component, the pneumatic circuit (<NUM>) including a set of interconnected pneumatic passages (<NUM>); and
a pressure-actuated switch (<NUM>) fluidly coupled to the pneumatic circuit (<NUM>),
wherein the pressure-actuated switch (<NUM>) is activated in response to a formation of a breach in the section of the component and an exposure of at least one of the pneumatic passages (<NUM>) of the pneumatic circuit (<NUM>) embedded in the section of the component;
wherein the pressure-actuated switch (<NUM>) is actuated in response to a reduction in pressure in the pneumatic circuit (<NUM>) due to the exposure of at least one of the pneumatic passages (<NUM>) of the pneumatic circuit (<NUM>) embedded in the section of the component,
characterised in that
the pressure-actuated switch (<NUM>) includes a fluid inlet (<NUM>) fluidly coupled to a pressurized source of cooling fluid (<NUM>, <NUM>) and a fluid outlet (<NUM>) fluidly coupled to the cooling circuit (<NUM>) of the component, and
wherein the activation of the pressure-actuated switch (<NUM>) increases the flow of cooling fluid (<NUM>, <NUM>) supplied to the cooling system (<NUM>) of the component.