Patent Description:
A gas turbine engine typically includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure compressor includes annular arrays ("rows") of stationary vanes that direct air entering the engine into downstream, rotating blades of the compressor. Collectively one row of compressor vanes and one row of compressor blades make up a "stage" of the compressor. Similarly, the high pressure turbine includes annular rows of stationary nozzle vanes that direct the gases exiting the combustor into downstream, rotating blades of the turbine. Collectively one row of nozzle vanes and one row of turbine blades make up a "stage" of the turbine. Typically, both the compressor and turbine include a plurality of successive stages.

Gas turbine engines, particularly aircraft engines, require a high degree of periodic maintenance. For example, periodic maintenance is often scheduled to allow internal components of the engine to be inspected for defects and subsequently repaired. Unfortunately, many conventional repair methods used for aircraft engines require that the engine be removed from the body of the aircraft and subsequently partially or fully disassembled. As such, these repair methods result in a significant increase in both the time and the costs associated with repairing internal engine components.

<CIT> (forming the basis for the preamble of the independent claim) discloses an apparatus for repairing turbine blades of a gas turbine engine by a laser welding operation, and a method using a miniaturized laser and related apparatus.

<CIT> discloses a maintenance device which includes a flexible member with an inspection end sized to be inserted through an inspection port of a workpiece such as a gas turbine engine or a blade of a gas turbine engine.

Accordingly, a system and method for performing an in situ repair of an internal component of a gas turbine engine would be welcomed within the technology.

Methods are generally provided for material build-up on a tip of a blade of a gas turbine engine. In one embodiment, the method includes inserting a material supply adjacent to the tip and directing a laser onto the interface of the material supply and the tip such that the material supply melts and attaches to the tip.

Methods are also generally provided for remotely stopping a crack in a component of a gas turbine engine. In one embodiment, the method includes inserting an integrated repair interface attached to a cable delivery system within a gas turbine engine; positioning the tip adjacent to a defect within a surface of the component; supplying a new material to the fillable area to fill the defect; and directing a laser to the new material within the fillable area to fuse the new material to the component within the defect.

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended Figs. , in which:.

In general, a system and method is provided for performing an in situ repair of an internal component of a gas turbine engine. According to the invention, the method includes include a repair tool configured to be inserted through an access port of the gas turbine engine to allow a repair tip or tip end of the tool to be positioned adjacent to a defect of an internal component of the engine, such as a crack, void, distressed area or any other defect defining a fillable volume. As will be described below, the repair tool may be configured to temporarily attach to the surface of the component, allowing precision work to be performed on the component. For example, the repair tool can supply a new material (solid or liquid) and direct a laser to heat to fuse new material within the crack to repair the defect. For example, if the new material is supplied as a solid, then the laser can heat and weld the material within the crack to repair the defect.

It should be appreciated that the disclosed system and method may generally be used to perform in situ repairs of internal components located within any suitable type of gas turbine engine, including aircraft-based turbine engines and land-based turbine engines, regardless of the engine's current assembly state (e.g., fully or partially assembled). Additionally, with reference to aircraft engines, it should be appreciated that the present subject matter may be implemented on-wing or off-wing.

Referring now to the drawings, <FIG> illustrates a cross-sectional view of one embodiment of a gas turbine engine <NUM> that may be utilized within an aircraft in accordance with aspects of the present subject matter, with the engine <NUM> being shown having a longitudinal or axial centerline axis <NUM> extending therethrough for reference purposes. In general, the engine <NUM> may include a core gas turbine engine (indicated generally by reference character <NUM>) and a fan section <NUM> positioned upstream thereof. The core engine <NUM> may generally include a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. In addition, the outer casing <NUM> may further enclose and support a booster compressor <NUM> for increasing the pressure of the air that enters the core engine <NUM> to a first pressure level. A high pressure, multi-stage, axial-flow compressor <NUM> may then receive the pressurized air from the booster compressor <NUM> and further increase the pressure of such air. The pressurized air exiting the high-pressure compressor <NUM> may then flow to a combustor <NUM> within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor <NUM>. The high energy combustion products are directed from the combustor <NUM> along the hot gas path of the engine <NUM> to a first (high pressure) turbine <NUM> for driving the high pressure compressor <NUM> via a first (high pressure) drive shaft <NUM>, and then to a second (low pressure) turbine <NUM> for driving the booster compressor <NUM> and fan section <NUM> via a second (low pressure) drive shaft <NUM> that is generally coaxial with first drive shaft <NUM>. After driving each of turbines <NUM> and <NUM>, the combustion products may be expelled from the core engine <NUM> via an exhaust nozzle <NUM> to provide propulsive jet thrust.

Additionally, as shown in <FIG>, the fan section <NUM> of the engine <NUM> may generally include a rotatable, axial-flow fan rotor assembly <NUM> that is configured to be surrounded by an annular fan casing <NUM>. It should be appreciated by those of ordinary skill in the art that the fan casing <NUM> may be configured to be supported relative to the core engine <NUM> by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes <NUM>. As such, the fan casing <NUM> may enclose the fan rotor assembly <NUM> and its corresponding fan rotor blades <NUM>. Moreover, a downstream section <NUM> of the fan casing <NUM> may extend over an outer portion of the core engine <NUM> so as to define a secondary, or by-pass, airflow conduit <NUM> that provides additional propulsive jet thrust.

It should be appreciated that, in several embodiments, the second (low pressure) drive shaft <NUM> may be directly coupled to the fan rotor assembly <NUM> to provide a direct-drive configuration. Alternatively, the second drive shaft <NUM> may be coupled to the fan rotor assembly <NUM> via a speed reduction device <NUM> (e.g., a reduction gear or gearbox) to provide an indirect-drive or geared drive configuration. Such a speed reduction device(s) may also be provided between any other suitable shafts and/or spools within the engine <NUM> as desired or required.

During operation of the engine <NUM>, it should be appreciated that an initial air flow (indicated by arrow <NUM>) may enter the engine <NUM> through an associated inlet <NUM> of the fan casing <NUM>. The air flow <NUM> then passes through the fan blades <NUM> and splits into a first compressed air flow (indicated by arrow <NUM>) that moves through conduit <NUM> and a second compressed air flow (indicated by arrow <NUM>) which enters the booster compressor <NUM>. The pressure of the second compressed air flow <NUM> is then increased and enters the high pressure compressor <NUM> (as indicated by arrow <NUM>). After mixing with fuel and being combusted within the combustor <NUM>, the combustion products <NUM> exit the combustor <NUM> and flow through the first turbine <NUM>. Thereafter, the combustion products <NUM> flow through the second turbine <NUM> and exit the exhaust nozzle <NUM> to provide thrust for the engine <NUM>.

The gas turbine engine <NUM> may also include a plurality of access ports defined through its casings and/or frames for providing access to the interior of the core engine <NUM>. For instance, as shown in <FIG>, the engine <NUM> may include a plurality of access ports <NUM> (only six of which are shown) defined through the outer casing <NUM> for providing internal access to one or both of the compressors <NUM>, <NUM> and/or for providing internal access to one or both of the turbines <NUM>, <NUM>. In several embodiments, the access ports <NUM> may be spaced apart axially along the core engine <NUM>. For instance, the access ports <NUM> may be spaced apart axially along each compressor <NUM>, <NUM> and/or each turbine <NUM>, <NUM> such that at least one access port <NUM> is located at each compressor stage and/or each turbine stage for providing access to the internal components located at such stage(s). In addition, the access ports <NUM> may also be spaced apart circumferentially around the core engine <NUM>. For instance, a plurality of access ports <NUM> may be spaced apart circumferentially around each compressor stage and/or turbine stage.

It should be appreciated that, although the access ports <NUM> are generally described herein with reference to providing internal access to one or both of the compressors <NUM>, <NUM> and/or for providing internal access to one or both of the turbines <NUM>, <NUM>, the gas turbine engine <NUM> may include access ports <NUM> providing access to any suitable internal location of the engine <NUM>, such as by including access ports <NUM> that provide access within the combustor <NUM> and/or any other suitable component of the engine <NUM>.

Referring now to <FIG>, a partial, cross-sectional view of the first (or high pressure) turbine <NUM> described above with reference to <FIG> is illustrated in accordance with embodiments of the present subject matter. As shown, the first turbine <NUM> may include a first stage turbine nozzle <NUM> and an annular array of rotating turbine blades <NUM> (one of which is shown) located immediately downstream of the nozzle <NUM>. The nozzle <NUM> may generally be defined by an annular flow channel that includes a plurality of radially-extending, circularly-spaced nozzle vanes <NUM> (one of which is shown). The vanes <NUM> may be supported between a number of arcuate outer bands <NUM> and arcuate inner bands <NUM>. Additionally, the circumferentially spaced turbine blades <NUM> may generally be configured to extend radially outwardly from a rotor disk (not shown) that rotates about the centerline axis <NUM> (<FIG>) of the engine <NUM>. Moreover, a turbine shroud <NUM> may be positioned immediately adjacent to the radially outer tips of the turbine blades <NUM> so as to define the outer radial flowpath boundary for the combustion products <NUM> flowing through the turbine <NUM> along the hot gas path of the engine <NUM>.

As indicated above, the turbine <NUM> may generally include any number of turbine stages, with each stage including an annular array of nozzle vanes and follow-up turbine blades <NUM>. For example, as shown in <FIG>, an annular array of nozzle vanes <NUM> of a second stage of the turbine <NUM> may be located immediately downstream of the turbine blades <NUM> of the first stage of the turbine <NUM>.

Moreover, as shown in <FIG>, a plurality of access ports <NUM> may be defined through the turbine casing and/or frame, with each access port <NUM> being configured to provide access to the interior of the turbine <NUM> at a different axial location. Specifically, as indicated above, the access ports <NUM> may, in several embodiments, be spaced apart axially such that each access port <NUM> is aligned with or otherwise provides interior access to a different stage of the turbine <NUM>. For instance, as shown in <FIG>, a first access port 62A may be defined through the turbine casing/frame to provide access to the first stage of the turbine <NUM> while a second access port 62B may be defined through the turbine casing/frame to provide access to the second stage of the turbine <NUM>.

It should be appreciated that similar access ports <NUM> may also be provided for any other stages of the turbine <NUM> and/or for any turbine stages of the second (or low pressure) turbine <NUM>. It should also be appreciated that, in addition to the axially spaced access ports <NUM> shown in <FIG>, access ports <NUM> may be also provided at differing circumferentially spaced locations. For instance, in one embodiment, a plurality of circumferentially spaced access ports may be defined through the turbine casing/frame at each turbine stage to provide interior access to the turbine <NUM> at multiple circumferential locations around the turbine stage.

Referring now to <FIG>, a partial, cross-sectional view of the high pressure compressor <NUM> described above with reference to <FIG> is illustrated in accordance with embodiments of the present subject matter. As shown, the compressor <NUM> may include a plurality of compressor stages, with each stage including both an annular array of fixed compressor vanes <NUM> (only one of which is shown for each stage) and an annular array of rotatable compressor blades <NUM> (only one of which is shown for each stage). Each row of compressor vanes <NUM> is generally configured to direct air flowing through the compressor <NUM> to the row of compressor blades <NUM> immediately downstream thereof.

Moreover, the compressor <NUM> may include a plurality of access ports <NUM> defined through the compressor casing/frame, with each access port <NUM> being configured to provide access to the interior of the compressor <NUM> at a different axial location. Specifically, in several embodiments, the access ports <NUM> may be spaced apart axially such that each access port <NUM> is aligned with or otherwise provides interior access to a different stage of the compressor <NUM>. For instance, as shown in <FIG>, first, second, third and fourth access ports 62a, 62b, 62c, 62d are illustrated that provide access to four successive stages, respectively, of the compressor <NUM>.

It should be appreciated that similar access ports <NUM> may also be provided for any of the other stages of the compressor <NUM> and/or for any of the stages of the low pressure compressor <NUM>. It should also be appreciated that, in addition to the axially spaced access ports <NUM> shown in <FIG>, access ports <NUM> may be also provided at differing circumferentially spaced locations. For instance, in one embodiment, a plurality of circumferentially spaced access ports may be defined through the compressor casing/frame at each compressor stage to provide interior access to the compressor <NUM> at multiple circumferential locations around the compressor stage.

Referring now to <FIG>, a simplified view of one embodiment of a system <NUM> for performing an in situ repair of an internal component of a gas turbine engine <NUM> are illustrated in accordance with aspects of the present subject matter. As shown, the system <NUM> may include a repair tool <NUM> configured to be inserted through an access port <NUM> of the gas turbine engine <NUM>, such as any of the access ports <NUM> described above with reference to <FIG>, to allow an in situ repair procedure to be performed on an internal component(s) (indicated by dashed lines <NUM>) of the engine <NUM>.

In general, the repair tool <NUM> may correspond to any suitable tool(s) and/or component(s) that may be inserted through an access port <NUM> of the gas turbine engine <NUM> and attach onto the surface <NUM> of the component <NUM> to perform precision work thereon. For example, an attachment mechanism <NUM> can temporarily attach onto the surface <NUM> so that the tool <NUM> can perform work at or near an identified defect <NUM> of the internal engine component(s) <NUM> being repaired (e.g., a turbine blade(s)). As such, the repair tool <NUM> may be temporarily attached to the surface <NUM> so as to allow for precision work at the defect <NUM> (e.g., with precision accuracy within about <NUM> or less, such as about <NUM> or less). As generically shown in <FIG>, a conduit <NUM> is attached to a working head <NUM> includes a work mechanism <NUM> controllable via a controller <NUM> (e.g., a computer or other programmable machine).

In one embodiment, the attachment mechanism <NUM> can be a tripod grip for a component <NUM> having a known shape and/or size. As shown in <FIG>, the component <NUM> is an airfoil tip <NUM> with a known shape and size (e.g., a nozzle and/or blade). In other embodiments, the component <NUM> can be a trailing edge and/or leading edge of the airfoil. The attachment mechanism <NUM> includes a plurality of grip arms <NUM> that attach the repair tool <NUM> onto the surface <NUM>. The grip arms <NUM> are brought together onto the edge of the tip <NUM> until the repair tool <NUM> is secured onto the tip <NUM>. In the embodiment shown, three grip arms <NUM> are included in the attachment mechanism <NUM>, although any suitable number of grip arms <NUM> may be utilized (e.g., three or more grip arms).

In another embodiment, the attachment mechanism <NUM> can be a suction cup attached onto the repair tool <NUM>. As shown in <FIG>, the attachment mechanism <NUM> includes a suction cup <NUM> that attach the repair tool <NUM> onto the surface <NUM>. In one embodiment, a vacuum can be applied within the suction cup <NUM> to hold the repair tool <NUM> onto the surface in place. The suction cup <NUM> can be constructed of a deformable, air-impervious material (e.g., a rubber material) that can form a suction attachment with the surface <NUM>. Although shown with one suction cup <NUM>, any number of suction cups can be utilized to secure the repair tool <NUM> onto the surface <NUM>. In yet another embodiment, an adhesive can be utilized to secure the repair tool <NUM> onto the surface <NUM>, such as a hot melt adhesive, epoxy material, etc. Then, the adhesive material can be melted to remove the repair tool <NUM> from the surface <NUM>.

Through the attachment mechanism <NUM>, the location of repair tool <NUM> can be precisely controlled and temporarily secured in place, which allows for precision work to be performed. In one embodiment, a working head <NUM> is positioned and secured adjacent to the identified defect <NUM> of the internal engine component(s) <NUM> being repaired (e.g., a turbine blade(s)). For example, as particularly shown in <FIG>, the defect <NUM> corresponds to a crack, void or other defective area formed along the exterior of the component <NUM> that defines an open or fillable volume <NUM> with a base <NUM> of the crack, void or other defective area.

As shown in <FIG>, the working head <NUM> includes a work mechanism <NUM> configured for addressing the defect <NUM>. In one embodiment, the new material can be supplied from a location exterior to the engine to the internal location of the defect <NUM> to allow the fillable volume <NUM> defined by the defect <NUM> to be filled with the new material. <FIG> shows the repair tool <NUM> configured to supply high velocity powder particles <NUM> from the exterior of the engine into the fillable volume <NUM> of the defect <NUM>. Upon impacting a surface of the defect <NUM>, the high velocity particles <NUM> may plastically deform and adhere to the surface, thereby filling-in the fillable volume <NUM> and repairing the defect <NUM>. For example, the particles can impact the surface within the defect <NUM> at a speed of about <NUM> meters per second (m/s) to about <NUM>/s.

The average size of the powder particles <NUM> can vary depending on their composition, gun type, nozzle type, gases used, etc. In most embodiments, the particle size and distribution can be about <NUM> to about <NUM> (e.g., about <NUM> to about <NUM> (i.e., <NUM> to about <NUM> mesh)). In certain embodiments, no more than about five percent of the particles are larger than about <NUM> (<NUM> mesh) and no more than about fifteen percent of the particles being smaller than about <NUM> (<NUM> mesh ).

The powder particles <NUM> can be supplied to the location of the defect via the repair tool <NUM> such that the fillable volume <NUM> may be filled-in with the powder particles <NUM>, thereby repairing the defect <NUM>. In several embodiments, the repair tool <NUM> may be configured to supply the powder particles <NUM> within the interior of the gas turbine engine <NUM>. For example, the powder particles <NUM> may be transported via the repair tool <NUM> from a location exterior to the gas turbine engine <NUM> to a location within the engine <NUM> to allow the powder particles <NUM> to be injected or otherwise directed into the fillable volume <NUM> defined by the defect <NUM>.

The particles <NUM> may be supplied via a carrier fluid (e.g., a carrier gas) that is inert to the coating deposition.

The powder particles <NUM> may then be heated to fuse the material within the fillable volume <NUM> to repair the defect <NUM>. For example, the repair tool <NUM> may include a laser <NUM> at its working end to heat the powder particles <NUM> during and/or after introduction into the defect <NUM>, thereby filling in the fillable volume <NUM> to bond the material within the defect <NUM>. For example, the working head <NUM> may include a laser <NUM> directed into the flow path of the powder particles <NUM> (e.g., into the defect <NUM>) to locally heat the base of the defect <NUM>, before, during, and/or after deposition of the new material (e.g., the powder particles <NUM>). For example, the laser <NUM> may direct electromagnetic radiation <NUM> having a wavelength of about <NUM> to about <NUM> into the defect <NUM> in the surface <NUM> of the component <NUM> at a power level of about <NUM> W to about 75kW. The electromagnetic radiation <NUM> can heat a precision weld within the base <NUM> of the defect <NUM> (e.g., at the deepest point from the surface <NUM> within the component <NUM>) to effectively stop the propagation of the defect <NUM> through the component <NUM>. <FIG> shows the laser <NUM> directing electromagnetic radiation <NUM> onto the new material within the defect <NUM>.

For example, the base <NUM> may be locally heated using the laser <NUM> to a temperature of about <NUM> to about <NUM> (e.g., about <NUM>° C to about <NUM>° C), particularly with the component <NUM> is constructed from a metal alloy or super-alloy such as a nickel-based alloy, a chromium-based alloy, etc..

According to the invention, the repair tool <NUM> includes one or more heating elements (indicated by dashed lines <NUM>) provided in operative association within the high temperature conduit <NUM>. As shown in the illustrated embodiment of <FIG>, the repair tool <NUM> may include a high temperature conduit <NUM> for transporting the metal particles from outside the engine <NUM> to the location of the defect <NUM>. Specifically, as shown in <FIG>, the high temperature conduit <NUM> may extend lengthwise between working head <NUM> located within the gas turbine engine <NUM> and a material supply end <NUM> located exterior to the engine <NUM>. The tip end of the tool <NUM> may generally be positioned adj acent to the location of the defect <NUM> for directing the particles <NUM> into the fillable volume <NUM>. Additionally, the material supply end <NUM> of the tool <NUM> may generally be configured to receive particles <NUM> from a particle source. For example, particles <NUM> contained within a chamber (or other suitable powder particle source) located exterior to the gas turbine engine <NUM> may be supplied to the material supply end <NUM> of the tool <NUM>. The particles <NUM> received at the material supply end <NUM> may then be directed through the high temperature conduit <NUM> to the tip end of the tool <NUM> to allow the metal particles to be delivered to the location of the defect <NUM>. It should be appreciated that the high temperature conduit <NUM> may generally be formed from any suitable high temperature material that allows the conduit <NUM> to serve as a fluid delivery means for the liquid metal. For example, in several embodiments, the high temperature conduit <NUM> may be formed from a ceramic material capable of withstanding temperatures above the melting temperature of the metal being supplied to the defect <NUM>. However, in other embodiments, the conduit <NUM> may be formed from any other suitable high temperature material.

According to the invention, the heating element(s) <NUM> are configured to generate heat within the high temperature conduit <NUM> as powder particles <NUM> is being supplied through the conduit <NUM> so as to allow for particle flow at the desired rate and speed. For example, in one embodiment, the heating element(s) <NUM> may correspond to a resisting heating element(s), such as one or more resistance wires, that is integrated into or incorporated within a wall(s) of the conduit <NUM>. However, in another embodiment, the heating element(s) <NUM> may correspond to any other suitable heat generating device(s) and/or component(s) that may be used to provide heating within the conduit <NUM> so as to maintain the temperature of the powder particles <NUM> at its desired delivery temperature. In one embodiment, the particles <NUM> are delivered to the defect <NUM> at a temperature within <NUM>% of its melting point (e.g., within <NUM>% of its melting point).

It should be appreciated that the powder particles <NUM> may be composed of any suitable metal material. For example, in one embodiment, the powder particles <NUM> may correspond to the parent metal material of the internal component <NUM> being repaired. In other embodiments, the powder particles <NUM> may correspond to any other metal material that is suitable for use as a repair material within a gas turbine engine <NUM>.

In one embodiment, the repair tool <NUM> includes an optical probe <NUM> adjacent to the working head <NUM> and configured to be used in association with the repair tool <NUM>. For instance, as shown in <FIG>, the optical probe <NUM> corresponds to a separate component configured to be used in combination with the repair tool <NUM> for repairing the defect <NUM>. However, in other embodiments, the optical probe <NUM> may be coupled to or integrated within the repair tool <NUM>. Additionally, as shown in <FIG>, the optical probe <NUM> has been inserted through the same access port <NUM> as the repair tool <NUM>. However, in other embodiments, the probe <NUM> may be inserted into a different access port <NUM> than the repair tool <NUM>, such as an access port <NUM> located adjacent to the access port <NUM> within which the repair tool <NUM> has been inserted.

In general, the optical probe <NUM> may correspond to any suitable optical device that allows images of the interior of the engine <NUM> to be captured or otherwise obtained. For instance, in several embodiments, the optical probe <NUM> may correspond to a borescope, videoscope, fiberscope or any other similar optical device known in the art that allows for the interior of a gas turbine engine <NUM> to be viewed through an access port <NUM>. In such embodiments, the optical probe <NUM> may include one or more optical elements (indicated schematically by dashed box <NUM>), such as one or more optical lenses, optical fibers, image capture devices, cables, and/or the like, for obtaining views or images of the interior of the engine <NUM> at a tip <NUM> of the probe <NUM> and for transmitting or relaying such images from the probe tip <NUM> along the length of the probe <NUM> to the exterior of the engine <NUM> for viewing by the personnel performing the repair procedure on the internal component(s) <NUM>. In addition, the probe <NUM> may include a light source (indicated by dashed box <NUM>) positioned at or adjacent to the probe tip <NUM> to provide lighting within the interior of the engine <NUM>.

As shown in <FIG> and <FIG>, the optical probe <NUM> may also include an articulation assembly <NUM> that allows the orientation of the probe tip <NUM> to be adjusted within the interior of the gas turbine engine <NUM>. For example, the articulation assembly <NUM> may allow for the probe tip <NUM> to be rotated or pivoted about a single axis or multiple axes to adjust the orientation of the tip <NUM> relative to the remainder of the probe <NUM>. It should be appreciated that the articulation assembly <NUM> may generally have any suitable configuration and/or may include any suitable components that allow for adjustment of the orientation of the probe tip <NUM> relative to the remainder of the probe <NUM>. For example, in one embodiment, a plurality of articulation cables <NUM> may be coupled between the probe tip <NUM> and one or more articulation motors <NUM>. In such an embodiment, by adjusting the tension of the cables <NUM> via the motor(s) <NUM>, the probe tip <NUM> may be reoriented within the gas turbine engine <NUM>.

In one particular embodiment, the articulation assembly <NUM> also controls the attachment mechanism <NUM> so as to temporarily attach to the surface <NUM> the component <NUM> in order to perform the desired work thereon.

Methods are generally provided for performing an in situ repair of an internal component of a gas turbine engine. In general, the methods are discussed herein with reference to the gas turbine engine <NUM> and the system <NUM> described above with reference to <FIG>. However, it should be appreciated by those of ordinary skill in the art that the disclosed methods may generally be implemented with gas turbine engines having any other suitable engine configuration and/or with systems having any other suitable system configuration. In addition, although the methods are discussed in a particular order for purposes of discussion, the methods discussed herein are not limited to any particular order or arrangement. One skilled in the art, using the disclosures provided herein, will appreciate that various steps of the methods disclosed herein can be omitted, rearranged, combined, and/or adapted in various ways without deviating from the scope of the present disclosure.

According to the invention, the method includes inserting a repair tool through an access port of the gas turbine engine such that the tool includes a tip end positioned within the engine; positioning the tip adjacent to a defect (e.g., a crack or other distress point) within the surface of the component; and temporarily attaching the tip adjacent to the defect to allow precision work to be performed. For example, as indicated above, the method may include positioning the tip end of the repair tool adjacent to a defect of an internal component of the gas turbine engine. As indicated above, the defect <NUM> may, for example, correspond to a crack, void or other defective area of an internal component <NUM> of the gas turbine engine <NUM>.

Moreover, the method may include performing precision repair work (e.g., supplying powder particles, heating, etc.) using the repair tool by temporarily attaching the tip end of the repair tool to the surface of the component.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any described methods.

Claim 1:
A method of remotely stopping a crack in a component (<NUM>) of a gas turbine engine (<NUM>), the method comprising:
inserting a repair tool (<NUM>) through an access port (<NUM>) of the gas turbine engine (<NUM>) such that the repair tool (<NUM>) includes a tip end (<NUM>) positioned within the gas turbine engine (<NUM>);
positioning the tip end (<NUM>) adjacent to a defect (<NUM>) within a surface of the component (<NUM>), wherein the defect (<NUM>) defines a fillable area;
temporarily attaching the tip end (<NUM>) adjacent to the defect (<NUM>) to allow precision work to be performed;
supplying a new material to the fillable area to fill the defect (<NUM>); and
directing a laser (<NUM>) to the new material within the fillable area to fuse the new material to the component (<NUM>) within the defect (<NUM>);
wherein the new material is supplied via a conduit (<NUM>) from a location exterior to the gas turbine engine (<NUM>);
the method being characterized in that:
the conduit (<NUM>) is a high temperature conduit (<NUM>) extending lengthwise between a material supply end of the repair tool (<NUM>) and the tip end (<NUM>) of the repair tool (<NUM>);
wherein one or more heating elements (<NUM>) are provided in operative association within the high temperature conduit (<NUM>);
and in that the method further comprises the step of generating heat by the one or more heating elements (<NUM>) within the high temperature conduit (<NUM>) as powder particles are supplied through the conduit (<NUM>) so as to allow for particle flow at the desired rate and speed.