Patent Publication Number: US-10316666-B2

Title: System and method for in situ balancing of a rotating component of a gas turbine engine

Description:
FIELD OF THE INVENTION 
     The present subject matter relates generally to gas turbine engines and, more particularly, to a system and method for performing an in situ repair of an internal rotating component of a gas turbine engine. More specifically, the present subject matter relates generally to in situ balancing of an internal rotating component of a gas turbine engine. 
     BACKGROUND OF THE INVENTION 
     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. 
     Gas turbine engines include various rotors in the typical form of bladed disks. Each rotor disk is specifically configured with a radially outer rim from which extends a row of blades. An axially thinner web extends radially inwardly from the rim and terminates in an axially thicker hub having a central bore therein. 
     A particular advantage of the bladed disk construction is that the integral disk may be smaller since no dovetails are used, and the blades are integrally formed around the disk rim. However, this construction increases repair difficulty since the blades are not readily individually removable from the disk. Minor repairs of the blade may be made in the bladed disk, but major repair thereof requires removal by cutting of corresponding portions of damaged blades or their complete removal, with the substitution thereof being made by welding or other metallurgical bonding process for achieving the original strength of the bladed disk. 
     An additional difficulty in the manufacture of the bladed disk is balancing thereof. All rotor components in a gas turbine engine must be suitably statically and dynamically balanced for minimizing rotary imbalance loads during operation for reducing vibration. The dovetail disk construction permits the rotor to be initially balanced during manufacture, with the individual blades being separately manufactured and matched in position on the disk for minimizing the resulting imbalance of the assembly thereof. 
     As such, a need exists for a method of in situ balancing of an internal rotating component, particularly a rotating disk, of a gas turbine engine. 
     BRIEF DESCRIPTION OF THE INVENTION 
     Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention. 
     Methods are generally provided for performing in situ balancing of an internal rotating component of a gas turbine engine. In one embodiment, the method includes inserting a repair tool through an access port of the gas turbine engine with the repair tool including a tip end positioned within the gas turbine engine and a material supply end positioned outside the gas turbine engine. The tip end of the repair tool is positioned adjacent to a surface of the internal rotating component of the gas turbine engine. A new material is supplied from the material supply end of the repair tool to the tip end of the repair tool; and is expelling from the tip end of the repair tool in a direction of the surface of the rotating component such that the new material is added onto a portion of the rotating part. 
     In another embodiment of the method, a repair tool is inserted through an access port of the gas turbine engine with the repair tool including a tip end positioned within the gas turbine engine and a material supply end positioned outside the gas turbine engine. The tip end of the repair tool is positioned adjacent to a surface of an internal component of the gas turbine engine. A solid filler material is supplied to the tip end of the repair tool, and is expelled from the tip end of the repair tool at a high flow velocity such that the solid filler material is directed onto the surface and adheres to the surface as the solid filler material impacts the internal component. 
     These and other features, aspects and advantages of the present invention will be better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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 figures, in which: 
         FIG. 1  illustrates a cross-sectional view of one embodiment of a gas turbine engine that may be utilized within an aircraft in accordance with aspects of the present subject matter; 
         FIG. 2  illustrates a partial, cross-sectional view of one embodiment of a turbine suitable for use within the gas turbine engine shown in  FIG. 1 , particularly illustrating access ports defined in the engine for providing internal access to the turbine; 
         FIG. 3  illustrates a partial, cross-sectional view of one embodiment of a compressor suitable for use within the gas turbine engine shown in  FIG. 1 , particularly illustrating access ports defined in the engine for providing internal access to the compressor; 
         FIG. 4  illustrates a simplified view of one embodiment of a system for performing an in situ balancing of an internal rotating component of a gas turbine engine in accordance with aspects of the present subject matter, particularly illustrating a repair tool inserted through an access port of the engine for supplying liquid metal to a defect of the internal component; 
         FIG. 5  illustrates a partial view of the repair tool and the internal component shown in  FIG. 4 , particularly illustrating a cross-sectional view of a portion of the repair tool; 
         FIG. 6  illustrates a flow diagram of one embodiment of a method for performing an in situ balancing of an internal rotating component of a gas turbine engine in accordance with aspects of the present subject matter; 
         FIG. 7  illustrates a simplified view of another embodiment of a system for performing an in situ balancing of an internal rotating component of a gas turbine engine in accordance with aspects of the present subject matter, particularly illustrating a repair tool inserted through an access port of the engine for supplying a high velocity solid filler material to a defect of the internal component; 
         FIG. 8  illustrates a partial view of the repair tool and the internal component shown in  FIG. 7 , particularly illustrating a cross-sectional view of a portion of the repair tool; 
         FIG. 9  illustrates a flow diagram of another embodiment of a method for performing an in situ balancing of an internal rotating component of a gas turbine engine in accordance with aspects of the present subject matter; and 
         FIG. 10  illustrates an exemplary bladed disk having new material added thereon for balancing thereof. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents. 
     In general, the present subject matter is directed to a system and method for performing in situ balancing (i.e., rebalancing) of an internal component of a gas turbine engine. Specifically, in several embodiments, the system may 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 surface of an internal component of the engine. As will be described below, the repair tool may be configured to supply a filler material from a location exterior to the engine to the surface of the rotating component to add additional new material to the component. For example, in one embodiment, the repair tool may be configured to supply liquid metal from the exterior of the engine onto the surface of the rotating component. The liquid metal may then cool and solidify onto the surface, thereby adding weight onto a portion of the surface of the rotating component. In another embodiment, the repair tool may be configured to supply high velocity solid filler material from the exterior of the engine onto the surface of the rotating component. Upon impacting a surface of the defect, the high velocity material may plastically deform and adhere to the surface, thereby adding weight to the surface of the rotating component. 
     It should be appreciated that the disclosed system and method may generally be used to perform in situ repairs (e.g., balancing) of internal rotating components (e.g., particularly rotating disks) located within any suitable type of gas turbine engine, including aircraft-based turbine engines and land-based turbine engines, regardless of the engine&#39;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. 
     At least some known rotor assemblies include components such as, but not limited to, disks, shafts, spools, bladed disks, seals, and/or bladed integrated and individual dovetail attached blades. A bladed disk is circumferentially continuous and has substantial hoop strength for withstanding the centrifugal loads developed by the blades as they rotate during operation about a longitudinal or axial centerline axis of the disk. The disk shape maximizes the strength thereof while minimizing undesirable weight for effectively supporting the blades over a substantial service life. 
     The rotor disks have various forms for supporting relatively large fan rotor blades and multiple rows of compressor blades decreasing in size for compressing air during operation. The air is mixed with fuel and ignited for generating hot combustion gases which flow downstream through various rows of turbine blades increasing in size on corresponding rotor disks therefor. In one particular configuration, the blades may be integrally formed with the rim of the disk in a unitary or one-piece construction. 
     When off-balance, the measured imbalance may be corrected by adding additional material diametrically oppositely from the angular position of the imbalance vector, such as at 180° for example. Material may be added to identified blades, or to the platform region between blades. Material may also be added to flanges on corresponding extension shafts of the bladed disk which are used for carrying torque load thereto from the low pressure turbine of the engine which powers the bladed disks. 
     As an exemplary rotor assembly,  FIG. 10  shows a rotating internal component  250  in the form of a bladed disk  252 . The exemplary bladed disk  252  includes a central disk thick section  254  and a plurality of compressor or turbine blades  256  that are prepared integrally with or metallurgically bonded to the thick section  254 . The bladed disk  252  may be made of any operable material, such as, for example, a nickel-based, cobalt-based, and/or iron-based superalloy. A platform  258  forms the radial outer area of the central disk thick section  254  between the plurality of compressor or turbine blades  256 . 
     The bladed disk may be included in any gas turbine engine, including propulsion engines such as turbofans, turboshafts, turboprops, etc. For example,  FIG. 1  illustrates a cross-sectional view of one embodiment of a gas turbine engine  10  that may be utilized within an aircraft in accordance with aspects of the present subject matter, with the engine  10  being shown having a longitudinal or axial centerline axis  12  extending therethrough for reference purposes. In general, the engine  10  may include a core gas turbine engine (indicated generally by reference character  14 ) and a fan section  16  positioned upstream thereof. The core engine  14  may generally include a substantially tubular outer casing  18  that defines an annular inlet  20 . In addition, the outer casing  18  may further enclose and support a booster compressor  22  for increasing the pressure of the air that enters the core engine  14  to a first pressure level. A high pressure, multi-stage, axial-flow compressor  24  may then receive the pressurized air from the booster compressor  22  and further increase the pressure of such air. The pressurized air exiting the high-pressure compressor  24  may then flow to a combustor  26  within which fuel is injected into the flow of pressurized air, with the resulting mixture being combusted within the combustor  26 . The high energy combustion products are directed from the combustor  26  along the hot gas path of the engine  10  to a first (high pressure) turbine  28  for driving the high pressure compressor  24  via a first (high pressure) drive shaft  30 , and then to a second (low pressure) turbine  32  for driving the booster compressor  22  and fan section  16  via a second (low pressure) drive shaft  34  that is generally coaxial with first drive shaft  30 . After driving each of turbines  28  and  32 , the combustion products may be expelled from the core engine  14  via an exhaust nozzle  36  to provide propulsive jet thrust. 
     Additionally, as shown in  FIG. 1 , the fan section  16  of the engine  10  may generally include a rotatable, axial-flow fan rotor assembly  38  that is configured to be surrounded by an annular fan casing  40 . It should be appreciated by those of ordinary skill in the art that the fan casing  40  may be configured to be supported relative to the core engine  14  by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes  42 . As such, the fan casing  40  may enclose the fan rotor assembly  38  and its corresponding fan rotor blades  44 . Moreover, a downstream section  46  of the fan casing  40  may extend over an outer portion of the core engine  14  so as to define a secondary, or by-pass, airflow conduit  48  that provides additional propulsive jet thrust. 
     It should be appreciated that, in several embodiments, the second (low pressure) drive shaft  34  may be directly coupled to the fan rotor assembly  38  to provide a direct-drive configuration. Alternatively, the second drive shaft  34  may be coupled to the fan rotor assembly  38  via a speed reduction device  37  (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  10  as desired or required. 
     During operation of the engine  10 , it should be appreciated that an initial air flow (indicated by arrow  50 ) may enter the engine  10  through an associated inlet  52  of the fan casing  40 . The air flow  50  then passes through the fan blades  44  and splits into a first compressed air flow (indicated by arrow  54 ) that moves through conduit  48  and a second compressed air flow (indicated by arrow  56 ) which enters the booster compressor  22 . The pressure of the second compressed air flow  56  is then increased and enters the high pressure compressor  24  (as indicated by arrow  58 ). After mixing with fuel and being combusted within the combustor  26 , the combustion products  60  exit the combustor  26  and flow through the first turbine  28 . Thereafter, the combustion products  60  flow through the second turbine  32  and exit the exhaust nozzle  36  to provide thrust for the engine  10 . 
     The gas turbine engine  10  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  14 . For instance, as shown in  FIG. 1 , the engine  10  may include a plurality of access ports  62  (only six of which are shown) defined through the outer casing  18  for providing internal access to one or both of the compressors  22 ,  24  and/or for providing internal access to one or both of the turbines  28 ,  32 . In several embodiments, the access ports  62  may be spaced apart axially along the core engine  14 . For instance, the access ports  62  may be spaced apart axially along each compressor  22 ,  24  and/or each turbine  28 ,  32  such that at least one access port  62  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  62  may also be spaced apart circumferentially around the core engine  14 . For instance, a plurality of access ports  62  may be spaced apart circumferentially around each compressor stage and/or turbine stage. 
     It should be appreciated that, although the access ports  62  are generally described herein with reference to providing internal access to one or both of the compressors  22 ,  24  and/or for providing internal access to one or both of the turbines  28 ,  32 , the gas turbine engine  10  may include access ports  62  providing access to any suitable internal location of the engine  10 , such as by including access ports  62  that provide access within the combustor  26  and/or any other suitable component of the engine  10 . 
     Referring now to  FIG. 2 , a partial, cross-sectional view of the first (or high pressure) turbine  28  described above with reference to  FIG. 1  is illustrated in accordance with embodiments of the present subject matter. As shown, the first turbine  28  may include a first stage turbine nozzle  66  and an annular array of rotating turbine blades  68  (one of which is shown) located immediately downstream of the nozzle  66 . The nozzle  66  may generally be defined by an annular flow channel that includes a plurality of radially-extending, circularly-spaced nozzle vanes  70  (one of which is shown). The vanes  70  may be supported between a number of arcuate outer bands  72  and arcuate inner bands  74 . Additionally, the circumferentially spaced turbine blades  68  may generally be configured to extend radially outwardly from a rotor disk (not shown) that rotates about the centerline axis  12  ( FIG. 1 ) of the engine  10 . Moreover, a turbine shroud  76  may be positioned immediately adjacent to the radially outer tips of the turbine blades  68  so as to define the outer radial flowpath boundary for the combustion products  60  flowing through the turbine  28  along the hot gas path of the engine  10 . 
     As indicated above, the turbine  28  may generally include any number of turbine stages, with each stage including an annular array of nozzle vanes and follow-up turbine blades  68 . For example, as shown in  FIG. 2 , an annular array of nozzle vanes  78  of a second stage of the turbine  28  may be located immediately downstream of the turbine blades  68  of the first stage of the turbine  28 . 
     Moreover, as shown in  FIG. 2 , a plurality of access ports  62  may be defined through the turbine casing and/or frame, with each access port  62  being configured to provide access to the interior of the turbine  28  at a different axial location. Specifically, as indicated above, the access ports  62  may, in several embodiments, be spaced apart axially such that each access port  62  is aligned with or otherwise provides interior access to a different stage of the turbine  28 . For instance, as shown in  FIG. 2 , a first access port  62 A may be defined through the turbine casing/frame to provide access to the first stage of the turbine  28  while a second access port  62 B may be defined through the turbine casing/frame to provide access to the second stage of the turbine  28 . 
     It should be appreciated that similar access ports  62  may also be provided for any other stages of the turbine  28  and/or for any turbine stages of the second (or low pressure) turbine  32 . It should also be appreciated that, in addition to the axially spaced access ports  62  shown in  FIG. 2 , access ports  62  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  28  at multiple circumferential locations around the turbine stage. 
     Referring now to  FIG. 3 , a partial, cross-sectional view of the high pressure compressor  24  described above with reference to  FIG. 1  is illustrated in accordance with embodiments of the present subject matter. As shown, the compressor  24  may include a plurality of compressor stages, with each stage including both an annular array of fixed compressor vanes  80  (only one of which is shown for each stage) and an annular array of rotatable compressor blades  82  (only one of which is shown for each stage). Each row of compressor vanes  80  is generally configured to direct air flowing through the compressor  24  to the row of compressor blades  82  immediately downstream thereof. 
     Moreover, the compressor  24  may include a plurality of access ports  62  defined through the compressor casing/frame, with each access port  62  being configured to provide access to the interior of the compressor  24  at a different axial location. Specifically, in several embodiments, the access ports  62  may be spaced apart axially such that each access port  62  is aligned with or otherwise provides interior access to a different stage of the compressor  24 . For instance, as shown in  FIG. 3 , first, second, third and fourth access ports  62   a ,  62   b ,  62   c ,  62   d  are illustrated that provide access to four successive stages, respectively, of the compressor  24 . 
     It should be appreciated that similar access ports  62  may also be provided for any of the other stages of the compressor  24  and/or for any of the stages of the low pressure compressor  22 . It should also be appreciated that, in addition to the axially spaced access ports  62  shown in  FIG. 3 , access ports  62  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  24  at multiple circumferential locations around the compressor stage. 
     Referring now to  FIGS. 4 and 5 , simplified views of one embodiment of a system  100  for performing an in situ balancing of an internal rotating component of a gas turbine engine  10  are illustrated in accordance with aspects of the present subject matter. As shown, the system  100  may include a repair tool  102  configured to be inserted through an access port  62  of the gas turbine engine  10 , such as any of the access ports  62  described above with reference to  FIGS. 1-3 , to allow an in situ repair procedure to be performed on an internal rotating component(s) (indicated by dashed lines  104 ) of the engine  10 . 
     In one embodiment, the repair tool  102  may correspond to any suitable tool(s) and/or component(s) that may be inserted through an access port  62  of the gas turbine engine  10  to allow a new material (e.g., a filler material, a new material, etc.) to be supplied within the engine  10  for adding material to a surface  105  of the rotating component  104  being repaired (e.g., a bladed disk). By supplying a filler material onto the surface  105  via the repair tool  102 , new material  108  may supply additional weight on a portion of the rotating component  104 , as shown in  FIG. 5 . 
     In several embodiments, the repair tool  102  may be configured to supply liquid metal within the interior of the gas turbine engine  10  as the filler material. For example, liquid metal may be transported via the repair tool  102  from a location exterior to the gas turbine engine  10  to a location within the engine  10  to allow the liquid metal to be coated or otherwise directed onto the surface  105  defined by the component  104 . The liquid metal may then solidify on the surface  105  as the metal cools. It should be appreciated that the liquid metal may generally correspond to any suitable metal material. For example, in one embodiment, the liquid metal may correspond to the parent metal material of the internal component  104  being repaired. In other embodiments, the liquid metal may correspond to any other metal material that is suitable for use as a repair material within a gas turbine engine  10 . 
     As shown in the illustrated embodiment, the repair tool  102  may include a high temperature conduit  110  for transporting the liquid metal from outside the engine  10  to the location of the defect  106 . Specifically, as shown in  FIG. 4 , the high temperature conduit  110  may extend lengthwise between a tip end  112  located within the gas turbine engine  10  and a material supply end  114  located exterior to the engine  10 . The tip end  112  of the tool  102  may generally be configured to be positioned adjacent to the component  104  for directing the liquid metal onto the surface  105 . Additionally, the material supply end  114  of the tool  102  may generally be configured to receive liquid metal from a liquid metal source. For example, as shown in  FIG. 5 , liquid metal contained within a furnace  116  (or other suitable liquid metal source) located exterior to the gas turbine engine  10  may be supplied to the material supply end  114  of the tool  102  (e.g., as indicated by arrow  118 ). The liquid metal received at the material supply end  114  may then be directed through the high temperature conduit  110  to the tip end  112  of the tool  102  to allow the liquid metal to be delivered to the surface  105 . 
     It should be appreciated that the high temperature conduit  110  may generally be formed from any suitable high temperature material that allows the conduit  110  to serve as a fluid delivery means for the liquid metal. For example, in several embodiments, the high temperature conduit  110  may be formed from a ceramic material capable of withstanding temperatures above the melting temperature of the metal being supplied onto the surface  105 . However, in other embodiments, the conduit  110  may be formed from any other suitable high temperature material. 
     Additionally, as particularly shown in  FIG. 5 , the repair tool  102  may include one or more heating elements (indicated by dashed lines  120 ) provided in operative association within the high temperature conduit  110 . In general, the heating element(s)  120  may be configured to generate heat within the high temperature conduit  110  as liquid metal is being supplied through the conduit  110  so as to maintain the metal in its liquid state. For example, in one embodiment, the heating element(s)  120  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  110 . However, in another embodiment, the heating element(s)  120  may correspond to any other suitable heat generating device(s) and/or component(s) that may be used to provide heating within the conduit  110  so as to maintain the temperature of the liquid metal at or above its melting temperature. 
     Moreover, in several embodiments, the repair tool  102  may also include a nozzle  122  positioned at or adjacent to the tip end  112  of the tool  102 . In general, the nozzle  122  may be configured to provide enhanced control of the direction of the flow of the liquid metal expelled from the tool  102 . For example, as shown in  FIG. 5 , the nozzle geometry may converge in the direction of the tip end  112  of the tool  102  to allow for precise control of the flow of the liquid metal relative to the surface  105 . It should be appreciated that, in one embodiment, the nozzle  122  may be formed integrally within the high temperature conduit  110 . Alternatively, the nozzle  122  may correspond to a separate component configured to be separately coupled to the conduit  110 . 
     Additionally, the system  100  may also include an optical probe  130  configured to be used in association with the repair tool  102 . For instance, as shown in  FIG. 4 , the optical probe  130  corresponds to a separate component configured to be used in combination with the repair tool  102  for adding new material  108  onto the surface  105 . However, in other embodiments, the optical probe  130  may be coupled to or integrated within the repair tool  102 . Additionally, as shown in  FIG. 4 , the optical probe  130  has been inserted through the same access port  62  as the repair tool  102 . However, in other embodiments, the probe  130  may be inserted into a different access port  62  than the repair tool  102 , such as an access port  62  located adjacent to the access port  62  within which the repair tool  102  has been inserted. 
     In general, the optical probe  130  may correspond to any suitable optical device that allows images of the interior of the engine  10  to be captured or otherwise obtained. For instance, in several embodiments, the optical probe  130  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  10  to be viewed through an access port  62 . In such embodiments, the optical probe  130  may include one or more optical elements (indicated schematically by dashed box  132 ), 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  10  at a tip  134  of the probe  130  and for transmitting or relaying such images from the probe tip  134  along the length of the probe  130  to the exterior of the engine  10  for viewing by the personnel performing the repair procedure on the internal component(s)  104 . In addition, the probe  130  may include a light source (indicated by dashed box  136 ) positioned at or adjacent to the probe tip  134  to provide lighting within the interior of the engine  10   
     As shown in  FIG. 4 , the optical probe  130  may also include an articulation assembly  138  that allows the orientation of the probe tip  134  to be adjusted within the interior of the gas turbine engine  10 . For example, the articulation assembly  138  may allow for the probe tip  134  to be rotated or pivoted about a single axis or multiple axes to adjust the orientation of the tip  134  relative to the remainder of the probe  130 . It should be appreciated that the articulation assembly  138  may generally have any suitable configuration and/or may include any suitable components that allow for adjustment of the orientation of the probe tip  134  relative to the remainder of the probe  130 . For example, in one embodiment, a plurality of articulation cables  140  may be coupled between the probe tip  134  and one or more articulation motors  142 . In such an embodiment, by adjusting the tension of the cables  140  via the motor(s)  142 , the probe tip  144  may be reoriented within the gas turbine engine  10 . 
     Referring now to  FIG. 6 , a flow diagram of one embodiment of a method  200  for performing an in situ repair of an internal component of a gas turbine engine is illustrated in accordance with aspects of the present subject matter. In general, the method  200  will be discussed herein with reference to the gas turbine engine  10  and the system  100  described above with reference to  FIGS. 1-5 . However, it should be appreciated by those of ordinary skill in the art that the disclosed method  200  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  FIG. 6  depicts steps performed in a particular order for purposes of illustration and 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. 
     As shown in  FIG. 6 , at ( 202 ), the method  200  may include 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 and a material supply end positioned outside the engine. For example, as indicated above, the repair tool  102  may include a high temperature conduit  110  extending lengthwise between its opposed ends  112 ,  114 . In such an embodiment, by inserting a portion of the conduit  110  through an access port  62  of the gas turbine engine  10 , the tip end  112  of the repair tool  102  may be located within the interior of the engine  10  while the material supply end  114  of the tool  102  may be positioned outside the engine  10 . Additionally, at ( 204 ), the method  200  may include positioning the tip end of the repair tool adjacent to a surface of an internal rotating component of the gas turbine engine. 
     Moreover, at ( 206 ), the method  200  may include supplying new material from the material supply end of the repair tool to the tip end of the repair tool. For example, as indicated above, the system  100  may include a new material source located exterior to the gas turbine engine  10 , such as a furnace  116  containing liquid metal. The new material may then be directed from the source  116  through the high temperature conduit  110  to the tip end  112  of the repair tool  102 . 
     Further, at ( 208 ), the method  200  may include expelling the new material from the tip end of the repair tool in a direction of the surface such that the new material is applied onto the component. Specifically, as indicated above, the liquid metal directed through the high temperature conduit  110  may be expelled from the tool  102  its tip end  112  and may flow onto the surface  105  of the component  104 . The liquid metal may then cool and solidify, thereby adding new material  108  onto the surface  105  of the component  104 . 
     Referring now to  FIGS. 7 and 8 , simplified views of another embodiment of a system  300  for performing an in situ repair of an internal component of a gas turbine engine  10  are illustrated in accordance with aspects of the present subject matter. As shown, the system  300  may include a repair tool  302  configured to be inserted through an access port  62  of the gas turbine engine  10 , such as any of the access ports  62  described above with reference to  FIGS. 1-3 , to allow an in situ balancing to be performed on an internal rotating component(s) (indicated by dashed lines  104 ) of the engine  10 . 
     Similar to the repair tool  102  described above, the repair tool  302  may be configured to be inserted through an access port  62  of the gas turbine engine  10  to allow a filler material to be supplied within the engine  10  for adding additional new material  108  onto the surface  105  of the internal rotating component(s)  104  to be repaired (e.g., a bladed disk(s)). However, unlike the embodiment described above, the filler material may correspond to a solid filler material (e.g., a solid powder material or a solid granularized material) configured to be directed onto the surface  105  at a high velocity such that the material adheres or mechanically bonds to the surface  105  ( FIG. 8 ) of the defect  106  as it impacts the internal component  104 . In such an embodiment, the solid filler material may be transported via the repair tool  302  from a location exterior to the gas turbine engine  10  to a location within the engine  10  to allow the material to be expelled or sprayed into the fillable volume  108  as a high velocity powder or projectile. Due to its high kinetic energy, the filler material may undergo plastic deformation and adhere to the surface  105  as the high velocity particles/projectiles impact the internal component  104 . 
     As shown in the illustrated embodiment, the repair tool  302  may include a supply conduit  310  for transporting the solid filler material from outside the engine  10  to the location of the surface  105 . Specifically, as shown in  FIG. 7 , the supply conduit  310  may extend lengthwise between a tip end  312  located within the gas turbine engine  10  and a material supply end  314  located exterior to the engine  10 . The tip end  312  of the repair tool  302  may generally be configured to be positioned adjacent to the location of the surface  105  for directing the filler material into the component  104 . Additionally, the material supply end  314  of the repair tool  302  may generally be in fluid communication with both a filler material source and a high pressure gas source. For example, as shown in  FIG. 7 , powder material (indicated by arrow  350 ) provided via a powder feeder or other powder source  352  may be mixed with a pressurized gas flow (indicated by arrow  354 ) received from a pressurized gas source  356 . The pressurized stream of powder/gas received at the material supply end  314  of the tool  302  may then be directed through the supply conduit  310  to the tip end  312  of the tool  302  for subsequent delivery to the surface  105 . 
     It should be appreciated that the solid filler material used within the system  300  may generally correspond to any suitable material that may be mechanically bonded to the inner surface  109  of the defect  106  via plastic deformation of the material upon impact with the internal component  104 , such as any suitable powder material or other material typically utilized within a cold spraying process. However, in several embodiments, the solid filler material may correspond to a metal-based solid powder material or a ceramic-based solid powder material. 
     It should also be appreciated that the gas mixed with the filler material may generally correspond to any suitable gas. However, in several embodiments, the gas may correspond to helium, nitrogen and/or air. In addition, in one embodiment, the gas flow provided from the pressurized gas source  356  may be heated. For example, the gas flow may be directed through a gas heater (not shown) positioned upstream of the location at which the gas flow is mixed with the solid filler material. 
     Additionally, the repair tool  302  may also include a nozzle  360  positioned at or adjacent to the tip end  312  of the repair tool  302  for increasing the flow velocity of the stream of filler material/gas being expelled or sprayed from the tool  302  into the surface  105 . As particularly shown in  FIG. 8 , the nozzle  360  may define a convergent-divergent geometry to allow the stream of filler material/gas to be accelerated as it flows through the nozzle  360 . For example, the nozzle may be configured as a De Laval nozzle and, thus, may include an upstream convergent section  362  and a downstream divergent section  364 . As such, as the stream of filler material/gas is directed from the convergent section  362  to the divergent section  364 , the stream may be accelerated to a significantly high velocity. The high velocity, high energy particles/projectiles expelled from the tip end  312  of the tool  302  may then impact the surface  105  and undergo plastic deformation, thereby allowing the particles/projectiles to mechanically bond to the surface  105 . 
     It should be appreciated that the nozzle  360  may generally be configured to accelerate the stream of filler material/gas to any suitable velocity that allows for the particles/projectiles to mechanically bond to the surface  105  upon impact with the internal component  104 . For example, in one embodiment, the nozzle  360  may be configured to accelerate the stream of filler material/gas to a supersonic flow velocity, such as a flow velocity greater than about 330 meters per second. 
     Additionally, as shown in  FIG. 7 , the system  300  may also include an optical probe  330  configured to be used in association with the repair tool  302 . In general, the optical probe  330  may be configured the same as or similar to the optical probe  130  described above with reference to  FIG. 4 . For example, the optical probe  330  may correspond to a borescope, videoscope or fiberscope or any other similar optical device known in the art that allows for the interior of a gas turbine engine  10  to be viewed through an access port  62 . In such an embodiment, the optical probe  330  may include one or more optical elements (indicated schematically by dashed box  332 ), 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  10  at a tip  334  of the probe  330  and for transmitting or relaying such images from the probe tip  334  along the length of the probe  330  to the exterior of the engine  10  for viewing by the personnel performing the repair procedure on the internal component(s)  104 . In addition, the probe  330  may include a light source (indicated by dashed box  336 ) positioned at or adjacent to the probe tip  334  to provide lighting within the interior of the engine  10  and an articulation assembly  338  (e.g., by including one or more articulation cables  340  and an associated articulation motor(s)  342 ) for adjusting the orientation of the probe tip  334  within the interior of the engine  10 . 
     Additionally, although not shown, it should be appreciated that the repair tool  302  may also include a suitable means for adjusting the orientation of its tip end  312  relative to the remainder of the tool  302 . For instance, the repair tool  302  may include an articulation assembly similar to the articulation assembly  338  used for the optical probe  330  to allow the location of the tip end  312  to be accurately positioned relative to the surface  105  being repaired. 
     Referring now to  FIG. 9 , a flow diagram of another embodiment of a method  400  for performing an in situ balancing of an internal rotating component of a gas turbine engine is illustrated in accordance with aspects of the present subject matter. In general, the method  400  will be discussed herein with reference to the gas turbine engine  10  described above with reference to  FIGS. 1-3  and the system  300  described above with reference to  FIGS. 7 and 8 . However, it should be appreciated by those of ordinary skill in the art that the disclosed method  400  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  FIG. 9  depicts steps performed in a particular order for purposes of illustration and 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. 
     As shown in  FIG. 9 , at ( 402 ), the method  400  may include 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 and a material supply end positioned outside the engine. For example, as indicated above, the repair tool  302  may include a high supply conduit  310  extending lengthwise between its opposed ends  312 ,  314 . In such an embodiment, by inserting a portion of the conduit  310  through an access port  62  of the gas turbine engine  10 , the tip end  312  of the repair tool  302  may be located within the interior of the engine  10  while the material supply end  314  of the tool  102  may be positioned outside the engine  10 . Additionally, at ( 304 ), the method  300  may include positioning the tip end of the repair tool adjacent to a surface of an internal rotating component of the gas turbine engine. 
     Moreover, at ( 406 ), the method  400  may include supplying a solid filler material to the tip end of the repair tool. For example, as indicated above, the repair tool  302  may be in fluid communication with both a pressurized gas source  356  and a filler material source  354  to allow a pressurized stream of filler material/gas to be received at the material supply end  314  of the tool  302 . The pressurized stream of filler material/gas may then be directed through the supply conduit  310  to the tip end  312  of the tool  302 . 
     Further, at ( 408 ), the method  400  may include expelling the solid filler material from the tip end of the repair tool at a high flow velocity such that the material is directed onto the surface and adheres to a surface of the defect as the material impacts the internal component. Specifically, as indicated above, the pressurized stream of filler material/gas may be directed through a nozzle  360  positioned at or adjacent to the tip end  312  of the tool  302  in order to accelerate the stream of filler material/gas to a substantially high flow velocity, such as a supersonic velocity. The high velocity, high energy particles/projectiles expelled from the tip end  312  of the tool  302  may then impact the surface  105  and undergo plastic deformation, thereby allowing the particles/projectiles to mechanically bond to the surface  105 . 
     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 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 are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.