Detectable data markers for measuring torsion in gas turbine engine components.

Abstract

This disclosure relates to a system (100) and method (200) for detecting twisting in a component (102) of a gas turbine engine (10). In one embodiment, the system (100) includes at least one detectable data marker (104) at least partially constructed within the component (102). Additionally, the system (100) includes a non-destructive inspection tool (106) configured to monitor the position of the detectable data marker (104) over time. Thus, twisting of the component (102) can be detected based on movement of the position of the detectable data marker (104).

Description

Technical Field

[0001] The present invention generally relates to gas turbine engines, and more specifically, to internally detectable data markers or reference points for measuring torsion in gas turbine engine components. Background Technology

[0002] A gas turbine engine generally comprises, in a continuous flow sequence, a compressor section, a combustion section, a turbine section, and an exhaust section. During operation, air enters the inlet of the compressor section, where one or more axial or centrifugal compressors progressively compress the air until it reaches the combustion section. Fuel mixes with the compressed air and burns within the combustion section to provide combustion gases. The combustion gases are then directed from the combustion section through a hot gas path defined within the turbine section, and finally exit from the turbine section via the exhaust section.

[0003] In a particular configuration, the turbine section comprises a high-pressure (HP) turbine and a low-pressure (LP) turbine in a continuous flow sequence. Each HP and LP turbine includes various rotatable turbine components, such as a rotor shaft, a rotor disk mounted on or otherwise supported by the rotor shaft, turbine blades mounted on the disk and extending radially from the disk's periphery, and various stationary turbine components, such as stator guide vanes or nozzles, turbine shrouds, and engine frames. The rotatable and stationary turbine components at least partially define the hot gas path through the turbine section. For example, gas turbine blades or impellers are generally airfoil-shaped and designed to convert the thermal and kinetic energy of the flowing gas into the mechanical rotation of the rotor. As the combustion gases flow through the hot gas path, thermal energy is transferred from the combustion gases to the rotatable and stationary turbine components. Such gas turbine engines are commonly used in aircraft.

[0004] During operation, various gas turbine engine components are susceptible to operating loads and / or environmental conditions that can cause dimensional distortion in the parts. While some distortion is acceptable, distortion exceeding a certain threshold may necessitate repair or replacement. Therefore, it is important to inspect such components for distortion during gas turbine engine operation. Traditional methods of component inspection require removing the parts from the gas turbine engine for further inspection, such as using a coordinate measuring machine (CMM) or a white light inspection system. However, such methods are time-consuming and expensive because they require engine shutdown and reduced operating time.

[0005] In light of the above, improved systems and methods for gas turbine engine components that do not require component removal would be advantageous. Summary of the Invention

[0006] Aspects and advantages of the invention will be set forth in part in the following description, and from this description may be apparent, or may be learned by practicing the invention.

[0007] On one hand, this disclosure relates to a system for detecting twist in components of a gas turbine engine. The system includes at least one detectable data marker at least partially constructed within the component. Additionally, the system includes a non-destructive inspection tool configured to monitor the position of the detectable data marker over time. Thus, twist in the component can be detected based on the movement of the position of the detectable data marker.

[0008] On the other hand, this disclosure relates to a method for detecting twisting in a component of a gas turbine engine. The method includes forming at least one detectable data marker at least partially constructed therein in the component of the gas turbine engine. Another step includes installing the component containing the detectable data marker within the gas turbine engine. The method also includes monitoring the position of the detectable data marker within the gas turbine engine component over time using a non-destructive inspection tool. Thus, twisting of the component can be detected based on the movement of the position of the detectable data marker.

[0009] Technical Solution 1. A system for detecting twisting in components of a gas turbine engine, the system comprising:

[0010] At least one detectable data marker is constructed, at least partially, within the component; and

[0011] A non-destructive inspection tool configured to monitor the location of the detectable data markers over time.

[0012] The twisting of the component is detected by the movement of the position of the detectable data marker.

[0013] Technical Solution 2. The system according to Technical Solution 1, characterized in that the component comprises a process-formed component at least partially constructed from a first material.

[0014] Technical Solution 3. The system according to Technical Solution 2, characterized in that the process-formed component is formed by at least one of composite layer lamination, pouring, molding or casting.

[0015] Technical Solution 4. The system according to Technical Solution 2, characterized in that the at least one detectable data marker is constructed from a second material, wherein the first material and the second material have different densities.

[0016] Technical Solution 5. The system according to Technical Solution 4, wherein the second material comprises at least one of the following: metal, metal alloy, radio frequency identification (RFID) chip, or polymer plastic.

[0017] Technical Solution 6. The system according to Technical Solution 5, wherein the metal or metal alloy comprises at least one of the following: tin, aluminum, nickel, lead, stainless steel, copper or phosphor bronze.

[0018] Technical Solution 7. The system according to Technical Solution 1, characterized in that the non-destructive inspection tool includes at least one of the following inspection methods: ultrasound, magnetic particles, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, or low coherence interference.

[0019] Technical Solution 8. The system according to Technical Solution 1, characterized in that the system further includes a plurality of detectable data markers located in a plurality of predetermined positions.

[0020] Technical Solution 9. The system according to Technical Solution 1, wherein the gas turbine engine comprises an aircraft engine.

[0021] Technical Solution 10. The system according to Technical Solution 9, characterized in that the aircraft engine remains installed on the aircraft during inspection.

[0022] Technical Solution 11. The system according to Technical Solution 1, characterized in that the components of the gas turbine engine include at least one of the following: turbine blades, turbine guide vanes, compressor blades, compressor guide vanes, fan blades, nozzles, discs, or shafts.

[0023] Technical Solution 12. A method for detecting torsion in components of a gas turbine engine, the method comprising:

[0024] The components of the gas turbine engine are formed with at least one detectable data marker that is at least partially constructed therein;

[0025] Install a component containing the detectable data marker within the gas turbine engine; and

[0026] The location of the detectable data markers within the gas turbine engine components over time is monitored using non-destructive inspection tools.

[0027] The twisting of the component is detected by the movement of the position of the detectable data marker.

[0028] Technical Solution 13. The method according to Technical Solution 12, characterized in that forming the component with at least one detectable data marker at least partially constructed therein further comprises forming the gas turbine engine component using a first material through at least one of composite laminate lamination, infusion, molding or casting.

[0029] Technical Solution 14. The method according to Technical Solution 13, characterized in that the at least one detectable data marker is constructed from a second material, wherein the first material and the second material have different densities.

[0030] Technical Solution 15. The method according to Technical Solution 14, wherein the second material comprises at least one of the following: metal, metal alloy, radio frequency identification (RFID) chip, or polymer plastic.

[0031] Technical Solution 16. The method according to Technical Solution 15, wherein the metal or metal alloy comprises at least one of the following: tin, aluminum, nickel, lead, stainless steel, copper or phosphor bronze.

[0032] Technical Solution 17. The method according to Technical Solution 12, characterized in that the non-destructive inspection tool includes at least one of the following inspection methods: ultrasound, magnetic particles, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, or low coherence interference.

[0033] Technical Solution 18. The method according to Technical Solution 12, characterized in that the method further includes forming a plurality of detectable data markers located at a plurality of predetermined positions on the components of the gas turbine engine.

[0034] Technical Solution 19. The method according to Technical Solution 11, wherein the gas turbine engine comprises an aircraft engine.

[0035] Technical Solution 20. The method according to Technical Solution 19, characterized in that the method further includes mounting the aircraft engine on the wing of the aircraft, and while the aircraft engine is running, monitoring the position of the detectable data marker within the component over time using the non-destructive inspection tool.

[0036] Technical Solution 21. A system (100) for detecting twist in a component (102) of a gas turbine engine (10), the system (100) comprising:

[0037] At least one detectable data marker (104) is at least partially constructed within the member (102); and

[0038] A non-destructive inspection tool (106) configured to monitor the position of the detectable data marker (104) over time.

[0039] The twisting of the component (102) is detected by the movement of the position of the detectable data marker (104).

[0040] Technical Solution 22. The system (100) according to Technical Solution 21 is characterized in that the component (102) includes a process-formed component (102) at least partially constructed of a first material.

[0041] Technical Solution 23. The system (100) according to Technical Solution 22 is characterized in that the process-formed component (102) is formed by at least one of composite layer lamination, pouring, molding or casting.

[0042] Technical Solution 24. The system (100) according to Technical Solution 22 is characterized in that the at least one detectable data marker (104) is constructed of a second material, wherein the first material and the second material have different densities.

[0043] Technical Solution 25. The system (100) according to Technical Solution 24 is characterized in that the second material comprises at least one of metal, metal alloy, radio frequency identification (RFID) chip or polymer plastic.

[0044] Technical Solution 26. The system (100) according to Technical Solution 25 is characterized in that the metal or metal alloy includes at least one of tin, aluminum, nickel, lead, stainless steel, copper or phosphor bronze.

[0045] Technical Solution 27. The system (100) according to Technical Solution 21 is characterized in that the non-destructive inspection tool (106) includes at least one of the following inspection methods (200): ultrasound, magnetic particles, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, or low coherence interference.

[0046] Technical Solution 28. The system (100) according to Technical Solution 21 is characterized in that the system (100) further includes a plurality of detectable data markers (104) located in a plurality of predetermined positions.

[0047] Technical Solution 29. The system (100) according to Technical Solution 21 is characterized in that the gas turbine engine (10) includes an aircraft engine, wherein the aircraft engine remains mounted on the aircraft during inspection.

[0048] Technical Solution 30. The system (100) according to Technical Solution 21 is characterized in that the components (102) of the gas turbine engine (10) include at least one of the following: turbine blades, turbine guide vanes, compressor blades, compressor guide vanes, fan blades, nozzles, discs or shafts.

[0049] Technical Solution 31. A method (200) for detecting twist in a component (102) of a gas turbine engine (10), the method (200) comprising:

[0050] The components (102) of the gas turbine engine (10) are formed with at least one detectable data marker (104) at least partially constructed therein;

[0051] The component (102) containing the detectable data marker (104) is installed within the gas turbine engine (10); and

[0052] The position of the detectable data marker (104) within the components (102) of the gas turbine engine (10) over time is monitored using a non-destructive inspection tool (106).

[0053] The twisting of the component (102) is detected by the movement of the position of the detectable data marker.

[0054] Technical Solution 32. The method (200) according to Technical Solution 31 is characterized in that forming the component (102) with at least one detectable data marker (104) at least partially constructed therein further comprises forming the gas turbine engine component (102) using a first material by at least one of composite laminate lamination, pouring, molding or casting.

[0055] Technical Solution 33. The method (200) according to Technical Solution 32 is characterized in that the at least one detectable data tag (104) is constructed of a second material, wherein the first material and the second material have different densities, wherein the second material includes at least one of the following: metal, metal alloy, radio frequency identification (RFID) chip or polymer plastic.

[0056] Technical Solution 34. The method (200) according to Technical Solution 31 is characterized in that the method (200) further includes forming a plurality of detectable data markers (104) located in a plurality of predetermined positions on a component (102) of the gas turbine engine (10).

[0057] Technical Solution 35. The method (200) according to Technical Solution 31, characterized in that the gas turbine engine (10) includes an aircraft engine, wherein the method further includes mounting the aircraft engine on the wing of an aircraft and, while the aircraft engine is in operation, monitoring the position of the detectable data marker (104) within the component (102) over time by means of the non-destructive inspection tool (106).

[0058] These and other features, aspects, and advantages of the invention will become more readily understood with reference to the following description and the 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 illustrate the principles of the invention. Attached Figure Description

[0059] The subject matter considered to be the invention is specifically pointed out and explicitly stated in the concluding section of the specification. However, the invention is best understood by referring to the following description, which is taken in conjunction with the accompanying drawings, wherein:

[0060] Figure 1 A cross-sectional view is shown of one embodiment of a gas turbine engine that can be used in an aircraft according to the present disclosure;

[0061] Figure 2 A schematic diagram of one embodiment of a system for detecting twisting in components of a gas turbine engine according to the present disclosure is shown;

[0062] Figure 3 Multiple images are shown generated by a non-destructive inspection tool of a system for detecting twisting in components of a gas turbine engine according to this disclosure; and

[0063] Figure 4 A flowchart illustrating an embodiment of a method for detecting twisting in components of a gas turbine engine according to the present disclosure is shown.

[0064] Parts list

[0065] 10 Gas Turbine Engine

[0066] 12 centerline axis

[0067] 14-core gas turbine engine

[0068] 16-fan section

[0069] 18 shell

[0070] 20 Circular Entrances

[0071] 22 booster compressor

[0072] 24 High-Pressure Compressor

[0073] 25 Gearbox

[0074] 26 burners

[0075] 28 High-Pressure Turbine

[0076] 30 High-pressure drive shaft

[0077] 32 Low-Pressure Turbine

[0078] 34 Low-pressure drive shaft

[0079] 36 exhaust nozzles

[0080] 37. Reduction Gear

[0081] 38 rotor assembly

[0082] 40 fan housing

[0083] 42 outlet guide vanes

[0084] 44 fan rotor blades

[0085] 46 Downstream Section

[0086] 48 airflow ducts

[0087] 50 airflow

[0088] 52 entrances

[0089] 54 airflow

[0090] 56 airflow

[0091] 58 airflow

[0092] 60 Combustion Products

[0093] 62A Proximity Port

[0094] 62B Proximity Port

[0095] 62 proximity ports

[0096] 66 Turbine Nozzle

[0097] 68 rotating turbine blades

[0098] 70 Nozzle Guide Vane

[0099] 72 arc-shaped outer band

[0100] 74 arc-shaped inner band

[0101] 76 Turbo Shield

[0102] 78 Nozzle Guide Vanes

[0103] 100 System

[0104] 102 components

[0105] 103rd floor

[0106] 104 detectable data tags

[0107] 105th floor

[0108] 106 Non-destructive Inspection Tools

[0109] 108 detector head

[0110] 110 controller

[0111] 112 User Interface

[0112] 200 methods

[0113] 202 Method Steps

[0114] 204 Method Steps

[0115] 206 Method Steps

[0116] 208 Method Steps

[0117] 210 Method and Steps

[0118] 212 Method and Steps. Detailed Implementation

[0119] Referring now to embodiments of the invention, one or more examples of the embodiments are illustrated in the figures. The various examples are provided in a manner that explains the invention rather than limits it. Indeed, it will be apparent to those skilled in the art that various modifications and variations can be made to the invention without departing from its scope or spirit. For example, a feature shown or described as part of one embodiment may be used in another embodiment to produce yet another embodiment. Therefore, it is intended that the invention cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0120] As used in this article, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another, without indicating the position or importance of a single component.

[0121] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction in which fluid flows out, while "downstream" refers to the direction in which fluid flows to.

[0122] Generally, this disclosure relates to a system and method for detecting twisting in components of a gas turbine engine. In one embodiment, the system includes at least one detectable data marker at least partially constructed within the component. Additionally, the system includes a non-destructive inspection tool configured to monitor the position of the detectable data marker over time. Thus, twisting of the component can be detected based on movement of the position of the detectable data marker.

[0123] This disclosure provides numerous advantages not found in the prior art. More specifically, by using a detectable reference material in gas turbine engine components, the system's non-destructive inspection tools can detect component distortion while the component is still mounted in the gas turbine engine or in operation. Consequently, the ability to eliminate the need for disassembly and on-site inspection reduces inspection costs and allows for real-time fault detection.

[0124] Now refer to the attached diagram, Figure 1 A cross-sectional view of one embodiment of a gas turbine engine 10 for use in an aircraft, according to various aspects of this subject matter, is shown, showing the engine 10 having a longitudinal or axial centerline axis 12 extending therethrough for reference purposes. Generally, the engine 10 may include a core gas turbine engine (generally indicated by 14) and a fan section 16 positioned upstream therefrom. The core engine 14 may generally include a substantially tubular casing 18 defining an annular inlet 20. Additionally, the casing 18 may further enclose and support a booster compressor 22 to increase the pressure of air entering the core engine 14 to a first pressure level. A high-pressure multistage axial flow compressor 24 may then receive pressurized air from the booster compressor 22 and further increase the pressure of this air. The pressurized air exiting the high-pressure compressor 24 may then flow to a combustor 26, where fuel is injected into the pressurized air stream, and the resulting mixture is burned within the combustor 26. High-energy combustion products are directed from combustor 26 along a hot gas path in engine 10 to a first (high-pressure) turbine 28 to drive high-pressure compressor 24 via a first (high-pressure) driveshaft 30, and then to a second (low-pressure) turbine 32 to drive supercharger compressor 22 and fan section 16 via a second (low-pressure) driveshaft 34, which is generally coaxial with the first driveshaft 30. After driving each of turbines 28 and 32, combustion products are discharged from core engine 14 through exhaust nozzles 36 to provide propulsive jet thrust.

[0125] In addition, such as Figure 1 As shown, the fan section 16 of engine 10 may generally include a rotatable axial flow fan rotor assembly 38, configured to be surrounded by an annular fan housing 40. Those skilled in the art will understand that the fan housing 40 may be configured to be supported relative to the core engine 14 by a plurality of generally radially extending, circumferentially spaced outlet guide vanes 42. Thus, the fan housing 40 may enclose the fan rotor assembly 38 and its corresponding fan rotor blades 44. Furthermore, a downstream section 46 of the fan housing 40 may extend over an external portion of the core engine 14 to define an auxiliary or bypass airflow duct 48, which would provide additional propulsive jet thrust.

[0126] It should be understood 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 reduction gear 37 (e.g., a reduction gear or gearbox) to provide an indirect-drive or gear-driven configuration. Such a reduction gear may also be provided between any other suitable shafts and / or axles within the engine 10 as desired or required.

[0127] During operation of engine 10, it should be understood that an initial airflow (indicated by arrow 50) enters engine 10 through the associated inlet 52 of fan housing 40. Airflow 50 is then directed through fan blades 44 and splits into a first compressed airflow (indicated by arrow 54) moving through duct 48 and a second compressed airflow (indicated by arrow 56) entering supercharger compressor 22. The pressure of the second compressed airflow 56 is then increased and it enters high-pressure compressor 24 (indicated by arrow 58). After mixing with and burning fuel in combustor 26, combustion products 60 exit combustor 26 and flow through first turbine 28. Thereafter, combustion products 60 flow through second turbine 32 and exit exhaust nozzle 36 to provide thrust to engine 10.

[0128] Now refer to Figure 2 This illustration shows a schematic diagram of one embodiment of a system 100 for detecting twist in a component 102 of a gas turbine engine 10. As used herein, twist generally refers to a change in the component 102 that indicates damage has occurred or may occur therein. Thus, the term twist generally includes non-material changes and more serious changes such as creep, flexure, bending, and / or deformation. Additionally, in some embodiments, as mentioned, the gas turbine engine 10 may include an aircraft engine. Thus, in such embodiments, the aircraft engine may be mounted on an aircraft during the inspection / checking procedures described herein. Therefore, the system 100 of this disclosure is configured to detect twist in the component 102 while the aircraft engine remains mounted on the aircraft. Additionally, the component 102 of the gas turbine engine 10 may include any engine component, including (but not limited to) turbine blades, turbine guide vanes, compressor blades, compressor guide vanes, fan blades, nozzles, discs, shafts, or any other suitable engine component.

[0129] More specifically, the component 102 of the gas turbine engine 10 may be a process-formed component, which is at least partially constructed of a first material, such as a metal, a metal alloy, or a polymer matrix composite. Additionally, in some embodiments, the process-formed component 102 may be formed by composite laminate lamination, infusion, molding, casting, or any other suitable process. For example, as... Figure 2 As shown in the diagram, component 102 is a composite layered component.

[0130] Therefore, as shown, system 100 also includes at least one detectable data marker 104, which is at least partially constructed within component 102 of gas turbine engine 10. More specifically, in the illustrated embodiment, during the stacking of layers 103, 105, the detectable data marker 104 can be easily placed between or near one or more layers 103, 105, such that the final component 102 has the marker 104 fully or partially embedded therein. In another embodiment, the detectable data marker 104 can be placed in the mold of component 102 during the manufacture of component 102, such that the marker 104 is integral with component 102 when installed in gas turbine engine 10. Additionally, as shown, system 100 may include a plurality of detectable data markers 104 located in a plurality of predetermined locations. For example, in the illustrated embodiment, four detectable data markers 104 are provided in component 102, but it should be understood that more than four or fewer detectable data markers 104 may be included at any location within or on component 102.

[0131] It should be understood that the detectable data tag 104 may be constructed from any suitable second material, for example, different from the first material of the component 102 of the gas turbine engine 10. Therefore, the first material of the component 102 and the second material of the detectable data tag 104 may have different or different densities. Consequently, the detectable data tag 104 can be easily observed by the non-destructive inspection tool 106, as will be described in more detail below. More specifically, in some embodiments, the second material of the detectable data tag 104 may be a metal, a metal alloy, a radio frequency identification (RFID) chip, a polymer plastic, or the like. Additionally, the metal or metal alloy may include tin, aluminum, nickel, lead, stainless steel, copper, or phosphor bronze.

[0132] Additionally, system 100 includes a non-destructive inspection tool 106 configured to monitor the position of the detectable data marker 104 over time. Thus, distortion of component 102 can be detected based on movement of the position of the detectable data marker 104. It should be understood that the non-destructive inspection tool 106 may include any suitable non-destructive inspection tool and / or method, including (but not limited to) X-ray, ultrasound, magnetic, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, low-coherence interferometry, or similar methods. For example, in one embodiment, the non-destructive inspection tool 106 may include a detector head 108 that allows the tool 106 to locate the detectable data marker 104 embedded within component 102. Additionally, the detector head 108 may optionally be coupled to a controller 110 and / or user interface 112 configured to generate images of component 102. Thus, the detector head 108 may be configured to acquire images of the detectable data marker 104 at a distance, or to acquire images by placing the head 108 near component 102. For example, as Figure 3 As shown in (1), the first image generated by the non-destructive inspection tool 106 shows the positions of the four data markers 104 (e.g., at the initial time), while Figure 3 (2) This shows a second image obtained by the non-destructive inspection tool 106, which depicts the twisted locations of the four data markers 104 (e.g., at some later time). Thus, the controller 110 can automatically determine the difference between the measurements of the first and second images to determine the total twist of the component 102. This process can also be performed manually by the user.

[0133] Now refer to Figure 4 A flowchart of one embodiment of a method 200 for detecting distortion in a component of a gas turbine engine 10 is shown. As shown at 202, method 200 includes forming a gas turbine engine component with at least one detectable data marker at least partially constructed therein. As shown at 204, the method includes mounting a component 102 containing a detectable data marker 104 within the gas turbine engine 10. As shown at 206, method 200 includes monitoring the position of the detectable data marker 104 within the component 102 over time using a non-destructive inspection tool 106. As shown at 208, if the position of one or more data markers 104 changes during monitoring, distortion 212 has occurred in the component 102. If the position of one or more data markers 104 remains unchanged during monitoring, distortion 210 has not yet occurred in the component 102.

[0134] In one embodiment, the step of forming a component 102 of a gas turbine engine 10 with at least one detectable data marker 104 at least partially constructed therein may include forming the component 102 using at least one of a first material through composite laminate lamination, infusion, molding, casting or other suitable processes and placing the marker 104 during the manufacturing process.

[0135] In another embodiment, method 200 may include a component 102 of a gas turbine engine 10 having a plurality of detectable data markers 104 located in a plurality of predetermined positions.

[0136] In an additional embodiment, method 200 may include mounting an aircraft engine on the wing of an aircraft and, while the aircraft engine is running, monitoring the location of detectable data markers within the gas turbine engine components over time using non-destructive inspection tools.

[0137] This written description uses examples to disclose the invention, including the best mode, and also enables any person skilled in the art to practice the invention, including making and using any device or system, and implementing any combination of methods. The patentable scope of the invention is defined by the claims and may include other examples that would 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 are not different from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

Claims

1. A system for detecting twist in a composite layered component of a gas turbine engine, the system comprising: At least one detectable data marker is constructed between the layers of the composite layered component; as well as A non-destructive inspection tool configured to monitor the location of the detectable data markers over time. The distortion of the composite layered member is detected by the movement of the position of at least one detectable data marker between the layers of the composite layered member and embedded within the composite layered member.

2. The system according to claim 1, characterized in that, The composite layered component includes a process-formed component that is at least partially constructed from a first material.

3. The system according to claim 2, characterized in that, The process-formed component is formed by at least one of composite laminate lamination, pouring, molding, or casting.

4. The system according to claim 2, characterized in that, The at least one detectable data marker is constructed from a second material, wherein the first material and the second material have different densities.

5. The system according to claim 4, characterized in that, The second material includes at least one of the following: metal, radio frequency identification (RFID) chip, or polymer plastic.

6. The system according to claim 5, characterized in that, The metal includes at least one of the following: tin, aluminum, nickel, lead, stainless steel, copper, or phosphor bronze.

7. The system according to claim 1, characterized in that, The non-destructive inspection tool includes at least one of the following inspection methods: ultrasound, magnetic particles, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, or low-coherence interference.

8. The system according to claim 1, characterized in that, The at least one detectable data tag includes multiple detectable data tags located at multiple predetermined locations.

9. The system according to claim 1, characterized in that, The gas turbine engine includes an aircraft engine.

10. The system according to claim 9, characterized in that, The aircraft engine remains installed on the aircraft during the inspection.

11. The system according to claim 1, characterized in that, The composite layered components of the gas turbine engine include at least one of the following: turbine blades, turbine guide vanes, compressor blades, compressor guide vanes, fan blades, nozzles, discs, or shafts.

12. The system according to claim 5, characterized in that, The metal includes metal alloys.

13. A method for detecting twist in a composite layered component of a gas turbine engine, the method comprising: The composite layered component of the gas turbine engine is formed with at least one detectable data marker constructed between the layers of the composite layered component; A composite layered component containing at least one detectable data marker is installed within the gas turbine engine; as well as The location of the at least one detectable data marker within the composite layered component over time is monitored using non-destructive inspection tools. The distortion of the composite layered member is detected by the movement of the position of at least one detectable data marker between the layers of the composite layered member and embedded within the composite layered member.

14. The method according to claim 13, characterized in that, Forming the composite layered member with at least one detectable data marker constructed between the layers of the composite layered member further includes forming the composite layered member using a first material through at least one of composite layer lamination, pouring, molding, or casting.

15. The method according to claim 14, characterized in that, The at least one detectable data marker is constructed from a second material, wherein the first material and the second material have different densities.

16. The method according to claim 15, characterized in that, The second material includes at least one of the following: metal, radio frequency identification (RFID) chip, or polymer plastic.

17. The method according to claim 16, characterized in that, The metal includes at least one of the following: tin, aluminum, nickel, lead, stainless steel, copper, or phosphor bronze.

18. The method according to claim 13, characterized in that, The non-destructive inspection tool includes at least one of the following inspection methods: ultrasound, magnetic particles, liquid penetrant, radiography, remote visual inspection (RVI), eddy current testing, or low-coherence interference.

19. The method according to claim 13, characterized in that, The method further includes forming the composite layered component of the gas turbine engine with at least one detectable data marker, which includes a plurality of detectable data markers located at a plurality of predetermined positions.

20. The method according to claim 13, characterized in that, The gas turbine engine includes an aircraft engine.

21. The method according to claim 20, characterized in that, The method further includes mounting the aircraft engine on the wing of the aircraft, and while the aircraft engine is running, monitoring the position of at least one detectable data marker between the layers of the composite layered component over time using the non-destructive inspection tool.

22. The method according to claim 16, characterized in that, The metal includes metal alloys.

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