Gas turbine engine with sensor assembly to detect torsional vibrations

Abstract

A gas turbine engine includes a fan section having a plurality of fan blades, the fan section configured to generate an airflow through the gas turbine engine; the airflow channel is provided with a core channel and a bypass channel separated from the core channel; a low speed shaft coupled to the plurality of fan blades and configured to rotate the plurality of fan blades; and a sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low speed shaft. The sensor assembly includes a plurality of dynamic pressure sensors in the airflow channel. A plurality of dynamic pressure sensors detect a dynamic pressure of the airflow passage, the dynamic pressure indicating torsional vibrations in the low speed shaft. The damping system is configured to damp torsional vibrations in the low speed shaft based on a correlation of the measured torsional vibrations and the experimented torsional vibrations.

Description

Technical Field

[0001] The present disclosure generally relates to a gas turbine engine having a sensor assembly for detecting torsional vibrations. Background Art

[0002] A gas turbine engine generally comprises a fan and a core section arranged in flow communication with each other, the core section comprising a low-pressure turbine. A low-speed shaft couples the low-pressure turbine to the fan. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] The foregoing and other features and advantages will be apparent from the following more particular description of various exemplary embodiments, as illustrated in the drawings, in which like reference numerals generally indicate identical, functionally similar, and / or structurally similar elements.

[0004] Figure 1 A schematic cross-sectional view of a non-ducted three-flow gas turbine engine according to the present disclosure, taken along the longitudinal centerline axis of the engine, is shown.

[0005] Figure 2 A schematic cross-sectional view of a ducted indirect drive gas turbine engine according to the present disclosure, taken along the longitudinal centerline axis of the engine, is shown.

[0006] Figure 3A A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0007] Figure 3B A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0008] Figure 3C A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0009] Figure 4A A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0010] Figure 4B A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0011] Figure 4C A method for using a gas turbine engine (such as for example Figure 2A first sensor assembly of a gas turbine engine Figure 4B A magnified view of .

[0012] Figure 4D A method for using a gas turbine engine (such as for example Figure 2 A second sensor assembly of a gas turbine engine Figure 4B An alternative magnified view of .

[0013] Figure 5 A method for calculating dynamic torque according to the present disclosure is shown.

[0014] Fig. 6A A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0015] Figure 6B and Figure 6C A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0016] Fig.6D and Fig. 6E A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0017] Figure 7 A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0018] Figure 8 The use of the present disclosure is shown Figure 7 A method for calculating dynamic torque of a sensor assembly.

[0019] Fig. 9 A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0020] Fig.10 The use of the present disclosure is shown Fig. 9 A method for calculating dynamic torque of a sensor assembly.

[0021] Fig.11A A schematic diagram of a sensor assembly and a shaft without torsional displacement is shown in accordance with the present disclosure.

[0022] Fig. 11B It shows that according to the present disclosure Fig.11ASchematic diagram of the sensor assembly and the shaft with torsional displacement.

[0023] Fig.12 A method for use with a gas turbine engine (such as for example with respect to Figure 2 A sensor assembly for a gas turbine engine described herein.

[0024] Fig.13 The use of the present disclosure is shown Fig.12 A method for calculating dynamic torque of a sensor assembly.

[0025] Fig.14A A schematic diagram of a sensor assembly and a shaft without torsional displacement is shown in accordance with the present disclosure.

[0026] Fig. 14B It shows that according to the present disclosure Fig.14A Schematic diagram of the sensor assembly and the shaft with torsional displacement.

[0027] Fig.15 A schematic diagram of components for taking corrective action in accordance with the present disclosure is shown. DETAILED DESCRIPTION

[0028] Additional features, advantages, and embodiments of the present disclosure are set forth or apparent by consideration of the following detailed description, drawings, and claims. Furthermore, the foregoing overview and the following detailed description of the present disclosure are exemplary and are intended to provide further explanation without limiting the scope of the present disclosure as claimed.

[0029] Various embodiments of the present disclosure are discussed in detail below. Although specific embodiments are discussed, this is for illustrative purposes only. Those skilled in the relevant art will recognize that other components and configurations may be used without departing from the spirit and scope of the present disclosure.

[0030] As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0031] The terms "upstream" and "downstream" refer to relative directions relative to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0032] The terms "front" and "rear" refer to relative positions within a turbine engine or vehicle, and refer to the normal operating attitude of the turbine engine or vehicle. For example, with respect to a turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

[0033] As used herein, the terms "low," "medium" (or "medium"), and "high," or their respective comparatives (e.g., "lower" and "higher," as applicable), when used with a compressor, turbine, shaft, fan, or turbine engine component, refer to relative pressures, relative speeds, relative temperatures, and / or relative power outputs within an engine, unless otherwise specified. For example, a "low power" setting defines an engine that is configured to operate at a power output lower than a "high power" setting for the engine, while a "medium power" setting defines an engine that is configured to operate at a power output higher than the "low power" setting and lower than the "high power" setting. The terms "low," "medium" (or "medium"), or "high" in such above-mentioned terms may additionally or alternatively be understood as relative to a minimum allowable speed, pressure, or temperature, or relative to a minimum or maximum allowable speed, pressure, or temperature for normal, desired, steady-state, etc., operation of the engine.

[0034] Unless otherwise indicated herein, the terms "coupled," "fixed," "attached," "connected," and the like refer to both direct coupling, fixing, attachment, or connection as well as indirect coupling, fixing, attachment, or connection through one or more intermediate components or features.

[0035] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0036] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of the turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of the turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of the turbine engine.

[0037] As used herein throughout the specification and claims, approximate language is applied to modify any quantitative representation that allows variation without causing a change in the basic function associated therewith. Therefore, the values ​​modified by terms such as "about," "approximately," "roughly," and "substantially" are not limited to the exact values ​​specified. In at least some cases, approximate language may correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct a component and / or system or to manufacture a component and / or system. For example, approximate language may refer to within a margin of one percent, two percent, four percent, ten percent, fifteen percent, or twenty percent of a single value, a range of values, and / or an endpoint of a limited range of values.

[0038] Here and throughout the specification and claims, range limitations are combined and interchanged. Unless the context or language indicates otherwise, such ranges are identified and include all subranges contained therein. For example, all ranges disclosed herein include the endpoints, and the endpoints can be combined independently of each other.

[0039] The present disclosure describes systems and methods for monitoring and detecting parameters that can indicate torsional dynamics (i.e., torsional vibration performance in an engine). In some examples, the systems and methods are used for experimental, demonstration, assembly or model engines to assist and enhance field monitoring of products (e.g., engines or aircraft). The systems and methods described herein use various sensors to monitor parameters that indicate torsional vibration. In some examples, the system includes a sensor that monitors pressure pulse information of a flow path. In some examples, the system includes a sensor that monitors motor parameters. In some examples, the system includes a sensor that uses a "magnetic pickup" to detect torsional displacement. In some examples, the system includes a rotary encoder. In some examples, the system includes one or more optical sensors or one or more high-speed cameras. The example systems and methods described herein can allow a correlation matrix for torsional performance to be developed based on engine testing.

[0040] Figure 1 A schematic diagram of a non-ducted, three-flow gas turbine engine 10 for an aircraft that may incorporate one or more embodiments of the present disclosure is shown. The gas turbine engine 10 is a "three-flow engine" because its architecture provides three different streams of thrust-producing airflow (labeled S1, S2, and S3) during operation, as described in further detail below.

[0041] like Figure 1 As shown, the gas turbine engine 10 defines an axial direction A, a radial direction R, and a circumferential direction C. In addition, the gas turbine engine 10 defines a longitudinal centerline axis 12 extending along the axial direction A. Generally, the axial direction A extends parallel to the longitudinal centerline axis 12, the radial direction R extends outwardly from the longitudinal centerline axis 12 and inwardly to the longitudinal centerline axis 12 in a direction perpendicular to the axial direction A, and the circumferential direction C extends three hundred and sixty degrees (360°) around the longitudinal centerline axis 12. The gas turbine engine 10 extends, for example, along the axial direction A between a forward end 14 and an aft end 16.

[0042] The gas turbine engine 10 includes a core engine 20 and a fan assembly 50 positioned upstream thereof. Generally, the core engine 20 includes a compressor section, a combustion section, a turbine section, and an exhaust section in a serial flow order. In particular, as shown in FIG. Figure 1As shown, the core engine 20 includes an engine core 18 and a core cover 22 that annularly surrounds the core engine 20. The engine core 18 includes a high pressure (HP) compressor 28, a combustor 30, and a high pressure (HP) turbine 32. The core engine 20 and the core cover 22 define a core inlet 24 having an annular shape. The core cover 22 further surrounds and supports a low pressure (LP) compressor 26 (also referred to as a supercharger), which is used to pressurize the air entering the core engine 20 through the core inlet 24. The HP compressor 28 receives the pressurized air from the LP compressor 26 and further increases the pressure of the air. The pressurized air flows downstream to the combustor 30, where fuel is injected into the pressurized air and ignited to increase the temperature and energy level of the pressurized air, thereby generating combustion gases.

[0043] The combustion gases flow downstream from the combustor 30 to the HP turbine 32. The HP turbine 32 drives the HP compressor 28 via a first shaft, also referred to as a high-pressure (HP) shaft 36 (also referred to as a high-speed shaft). In this regard, the HP turbine 32 is drivingly coupled to the HP compressor 28. The combustion gases then flow to a power turbine or low-pressure (LP) turbine 34. The LP turbine 34 drives the LP compressor 26 and components of the fan assembly 50 via a second shaft, also referred to as a low-pressure (LP) shaft 38 (also referred to as a low-speed shaft). In this regard, the LP turbine 34 is drivingly coupled to the LP compressor 26 and components of the fan assembly 50. Figure 1 In the embodiment of FIG. 1 , the low speed shaft 38 is coaxial with the high speed shaft 36. After driving each of the HP turbine 32 and the LP turbine 34, the combustion gases exit the engine core 18 through a core exhaust nozzle 40. The core engine 20 defines a core flow path, also referred to as a core duct 42, extending between the core inlet 24 and the core exhaust nozzle 40. The core duct 42 is an annular duct positioned generally inwardly of the core shroud 22 in the radial direction R.

[0044] The fan assembly 50 includes a primary fan 52. Figure 1 In the preferred embodiment, the primary fan 52 is an open rotor fan, also referred to as a non-ducted fan. However, in other embodiments, the primary fan 52 may be ducted, for example, by a fan housing or nacelle circumferentially surrounding the primary fan 52. The primary fan 52 includes an array of fan blades 54 (in Figure 1 The fan blades 54 are rotatably coupled to the hub 53 for rotation about the longitudinal centerline axis 12 via the fan shaft 56. Figure 1 As shown, for example, in an indirect drive configuration, the fan shaft 56 is coupled to the low speed shaft 38 via a reduction gearbox (also referred to as a gearbox assembly 55). The gearbox assembly 55 is Figure 1Schematically shown in FIG. The gearbox assembly 55 includes a plurality of gears for adjusting the rotational speed of the fan shaft 56, and therefore the rotational speed of the primary fan 52 relative to the low speed shaft 38, to a more efficient rotating fan speed. The fan gearbox assembly can have a gear ratio of 4:1 to 12:1, or 7:1 to 12:1, or 4:1 to 10:1, or 5:1 to 9:1, or 6:1 to 9:1, and can be constructed in an epicyclic star gear or planetary gear configuration. The gearbox can be a single stage or compound gearbox.

[0045] The fan blades 54 may be arranged equally spaced about the longitudinal centerline axis 12. Each fan blade 54 has a root and a tip and a span defined therebetween. Each fan blade 54 defines a central blade axis 57. Figure 1 In some embodiments, each fan blade 54 of the primary fan 52 can rotate about its respective central blade axis 57, for example in unison with each other. One or more actuators 58 are controlled to pitch the fan blades 54 about their respective central blade axes 57. In other embodiments, each fan blade 54 is fixed or cannot pitch about the central blade axis 57.

[0046] The fan assembly 50 also includes a fan guide vane array 60 including fan guide vanes 62 ( Figure 1 Only one is shown in Figure 1). Figure 1 In the embodiment of the present invention, the fan guide vanes 62 are not rotatable about the longitudinal centerline axis 12. Each fan guide vane 62 has a root and a tip, and a span defined therebetween. The fan guide vanes 62 may be configured as follows: Figure 1 The fan guide vanes 62 are shown unshielded, or may be shielded, for example, by an annular shroud spaced outwardly from the tips of the fan guide vanes 62 in the radial direction R. Each fan guide vane 62 defines a central vane axis 64. Figure 1 In some embodiments, each fan guide vane 62 of the fan guide vane array 60 can rotate about its respective central vane axis 64, for example in unison with each other. One or more actuators 66 are controlled to pitch the fan guide vanes 62 about their respective central vane axis 64. In other embodiments, each fan guide vane 62 is fixed or cannot pitch about the central vane axis 64. The fan guide vanes 62 are mounted to the fan housing 70.

[0047] The fan cowl 70 annularly surrounds at least a portion of the core cowl 22 and is positioned generally outside the core cowl 22 in the radial direction R. In particular, a downstream section of the fan cowl 70 extends over a front portion of the core cowl 22 to define a fan flow path, also referred to as a fan duct 72. Incoming air enters the fan duct 72 through a fan duct inlet 76 and exits through a fan exhaust nozzle 78 to generate propulsive thrust. The fan duct 72 is an annular duct positioned generally outside the core duct 42 in the radial direction R. The fan cowl 70 and the core cowl 22 are connected together and are connected by a plurality of struts 74 ( Figure 1 The plurality of struts 74 are shaped aerodynamically to guide the air flowing therethrough. Other struts besides the plurality of struts 74 may be used to connect and support the fan shroud 70 and / or the core shroud 22.

[0048] The gas turbine engine 10 further defines or includes an inlet duct 80. The inlet duct 80 extends between the engine inlet 82 and the core inlet 24 and the fan duct inlet 76. The engine inlet 82 is generally defined at the front end of the fan casing 70 and is positioned between the primary fan 52 and the fan guide vane array 60 along the axial direction A. The inlet duct 80 is an annular duct positioned inside the fan casing 70 along the radial direction R. Air flowing downstream along the inlet duct 80 is divided (not necessarily evenly) into the core duct 42 and the fan duct 72 by the flow divider 84 of the core casing 22. The inlet duct 80 is wider than the core duct 42 along the radial direction R. The inlet duct 80 is also wider than the fan duct 72 along the radial direction R.

[0049] The fan assembly 50 also includes an intermediate fan 86. The intermediate fan 86 includes a plurality of intermediate fan blades 88 ( Figure 1 Only one is shown in FIG. 1 ). A plurality of intermediate fan blades 88 are rotatable, for example, about the longitudinal centerline axis 12. The intermediate fan 86 is drivingly coupled to the LP turbine 34 via the low-speed shaft 38. A plurality of intermediate fan blades 88 may be arranged at equal circumferential intervals about the longitudinal centerline axis 12. A plurality of intermediate fan blades 88 are annularly surrounded (e.g., ducted) by the fan shroud 70. In this regard, the intermediate fan 86 is positioned inside the fan shroud 70 along the radial direction R. The intermediate fan 86 is positioned within the inlet duct 80 upstream of the core duct 42 and the fan duct 72. The ratio of the span of the fan blade 54 to the span of the intermediate fan blade 88 (the span is measured from the root to the tip of the corresponding blade) is greater than two and less than ten to achieve the desired benefits of the third flow S3, in particular the additional thrust provided to the gas turbine engine 10 by the ratio, which may enable smaller diameter fan blades 54, which in turn may be beneficial to engine installation.

[0050] Thus, the air flowing through the inlet duct 80 flows through the plurality of intermediate fan blades 88 and is accelerated downstream thereof. At least a portion of the air accelerated by the intermediate fan blades 88 flows into the fan duct 72 and is ultimately discharged through the fan exhaust nozzle 78 to generate propulsive thrust. In addition, at least a portion of the air accelerated by the plurality of intermediate fan blades 88 flows into the core duct 42 and is ultimately discharged through the core exhaust nozzle 40 to generate propulsive thrust. Typically, the intermediate fan 86 is a compression device positioned downstream of the engine inlet 82. The intermediate fan 86 is operable to accelerate air entering the fan duct 72, also referred to as a secondary bypass passage.

[0051] During operation of the gas turbine engine 10, the initial or incoming airflow passes through the fan blades 54 of the primary fan 52 and is divided into a first airflow and a second airflow. The first airflow bypasses the engine inlet 82 and flows in the radial direction R generally in the axial direction A outside the fan casing 70. The first airflow accelerated by the fan blades 54 passes through the fan guide vanes 62 and thereafter continues downstream to generate a primary propulsive flow or a first thrust flow S1. Most of the net thrust generated by the gas turbine engine 10 is generated by the first thrust flow S1. The second airflow enters the inlet duct 80 through the engine inlet 82.

[0052] The second airflow flowing downstream through the inlet duct 80 flows through a plurality of intermediate fan blades 88 of the intermediate fan 86 and is thereby compressed. The second airflow flowing downstream of the intermediate fan blades 88 is split by the splitter 84 located at the front end of the core cover 22. Specifically, a portion of the second airflow flowing downstream of the intermediate fan 86 flows into the core duct 42 through the core inlet 24. The portion of the second airflow flowing into the core duct 42 is gradually compressed by the LP compressor 26 and the HP compressor 28, and is ultimately discharged to the combustion section. The discharged pressurized air flow flows downstream to the combustor 30, where fuel is introduced to generate combustion gases or products.

[0053] The combustor 30 defines an annular combustion chamber that is substantially coaxial with the longitudinal centerline axis 12. The combustor 30 receives pressurized air from the HP compressor 28 via a pressure compressor discharge outlet. A portion of the pressurized air flows into the mixer. Fuel is injected through a fuel nozzle (omitted for clarity) to mix with the pressurized air, thereby forming a fuel-air mixture provided to the combustion chamber for combustion. Ignition of the fuel-air mixture is completed by one or more igniters (omitted for clarity), and the resulting combustion gases flow along an axial direction A toward and into the first-stage turbine nozzle of the HP turbine 32. The first-stage turbine nozzle 33 is defined by an annular flow channel including a plurality of radially extending, circumferentially spaced nozzle vanes 35 that rotate the combustion gases so that the combustion gases flow at an angle and impact the first-stage turbine blades of the HP turbine 32. The combustion gases leave the HP turbine 32 and flow through the LP turbine 34, and leave the core duct 42 through the core exhaust nozzle 40 to produce a core air flow, also referred to as a second thrust flow S2. As described above, the HP turbine 32 drives the HP compressor 28 via the high-speed shaft 36 , and the LP turbine 34 drives the LP compressor 26 , the primary fan 52 , and the intermediate fan 86 via the low-speed shaft 38 .

[0054] Another portion of the second airflow flowing downstream of the intermediate fan 86 is split by the splitter 84 into the fan duct 72. The air enters the fan duct 72 through the fan duct inlet 76. The air flows through the fan duct 72 generally in the axial direction A, and is eventually discharged from the fan duct 72 through the fan exhaust nozzle 78 to generate a third flow, also referred to as a third thrust flow S3.

[0055] The third thrust flow S3 is a secondary air flow that increases fluid energy to produce a small portion of the total propulsion system thrust. In some embodiments, the pressure ratio of the third flow is higher than the pressure ratio of the primary propulsion flow (e.g., bypass or propeller-driven propulsion flow). Thrust can be generated by a dedicated nozzle or by mixing the secondary air flow with the primary propulsion flow or core air flow into, for example, a common nozzle. In some embodiments, the operating temperature of the secondary air flow is lower than the maximum compressor discharge temperature of the engine. In addition, in some embodiments, aspects of the third flow (e.g., air flow characteristics, mixing characteristics, or exhaust characteristics) and the percentage contribution to the total thrust are passively adjusted during engine operation or can be purposefully modified by using engine control features (e.g., fuel flow, motor power, variable stator, variable inlet guide vane, valve, variable exhaust geometry, or fluid characteristics) to adjust or improve overall system performance under a wide range of potential operating conditions.

[0056] Figure 1The gas turbine engine 10 depicted in the figure is only an example. In other embodiments, the gas turbine engine 10 can have any other suitable configuration. For example, in other embodiments, the primary fan 52 can be constructed in any other suitable manner (e.g., as a fixed pitch fan), and can also be supported using any other suitable fan frame (e.g., fan frame 98 of Figure 3) configuration. In other embodiments, the primary fan 52 can be ducted through a fan case or a nacelle so that a bypass passage is defined between the fan case and the fan cover 70. In addition, in other embodiments, any other suitable number or configuration of compressors, turbines, shafts, or combinations thereof can be provided. In other embodiments, aspects of the present disclosure can be incorporated into any other suitable turbine engine (e.g., a turbofan engine, a propeller fan engine, a turbojet engine, a turboprop engine, a turboshaft engine, and / or a turbine engine defining two flows (e.g., a bypass flow and a core air flow)).

[0057] In addition, for Figure 1 In the depicted embodiment, the gas turbine engine 10 includes an electric machine 90, which may be an electric generator, operatively coupled to its rotating components. In this regard, the gas turbine engine 10 is a hybrid electric propulsion machine. In particular, as Figure 1 As shown, the motor 90 is operably coupled to the low speed shaft 38. The motor 90 may be coupled directly or, for example, via a gearbox assembly 92 (in Figure 1 ) is indirectly mechanically connected to the low-speed shaft 38. In addition, although in this embodiment, the motor 90 is operably coupled to the low-speed shaft 38 at the rear end of the low-speed shaft 38, the motor 90 may be coupled to the low-speed shaft 38 at any suitable location, or may be coupled to other rotating components of the gas turbine engine 10, such as the high-speed shaft 36 or the low-speed shaft 38. For example, in some embodiments, the motor 90 may be coupled to the low-speed shaft 38 and positioned in front of the intermediate fan 86 in the axial direction.

[0058] In some embodiments, motor 90 may be an electric motor that is operable to drive or propel low-speed shaft 38, for example, during engine bursts. In other embodiments, motor 90 may be a generator that is operable to convert mechanical energy into electrical energy. In this manner, the electricity generated by motor 90 may be directed to various engine and / or aircraft systems. In some embodiments, motor 90 may be a dual-function motor / generator. Motor 90 includes rotor 94 and stator 96. Rotor 94 is coupled to low-speed shaft 38 and rotates as low-speed shaft 38 rotates. In this manner, rotor 94 rotates relative to stator 96, thereby generating electricity. Although in Figure 1The motor 90 has been described and shown as having a particular configuration, but the present disclosure may be applied to motors having alternative configurations. For example, the rotor 94 and / or the stator 96 may have a different configuration, or may be configured in a manner similar to that of the rotor 94 and / or the stator 96. Figure 1 Different arrangements are shown.

[0059] Figure 2 A schematic cross-sectional view of a ducted indirect drive gas turbine engine 100 (also referred to as turbine engine 100) according to an embodiment of the present disclosure is shown, taken along a longitudinal centerline axis 112 of the gas turbine engine 100. The gas turbine engine 100 (also referred to herein as turbine engine 100) includes, in downstream serial flow relationship: a fan section 114 including a fan 102; a compressor section 116 including a supercharger or low pressure (LP) compressor 121 and a high pressure (HP) compressor 118; a combustion section 128 including a combustor 130; a turbine section 133 including an HP turbine 134, an LP turbine 120; and an exhaust nozzle 138.

[0060] The fan section 114 includes a fan case 140 fixed to the nacelle (omitted for clarity), surrounding the fan 102. The fan 102 includes a plurality of fan blades 124 radially arranged around the longitudinal centerline axis 112. The HP compressor 118, the combustor 130, and the HP turbine 134 form an engine core 144 of the turbine engine 100, which generates combustion gases. The engine core 144 is surrounded by a core case 131 coupled to the fan case 140. The fan case 140 is supported relative to the turbine by circumferentially spaced outlet guide vanes 182.

[0061] A high-speed shaft 148 (also referred to herein as a high-pressure shaft) is coaxially disposed about the longitudinal centerline axis 112 of the turbine engine 100 and drivingly connects the HP turbine 134 to the HP compressor 118. A low-speed shaft 122 (also referred to herein as a low-pressure shaft 122) coaxially disposed about the longitudinal centerline axis 112 of the turbine engine 100 and within the larger diameter annular high-speed shaft 148 drives the LP turbine 120 to the LP compressor 121 and the fan 102 (either directly or through a gearbox assembly 150). The high-speed shaft 148 and the low-speed shaft 122 are rotatable about the longitudinal centerline axis 112.

[0062] The LP compressor 121 and the HP compressor 118 include respective multiple compressor stages 152, 154, wherein respective sets of compressor blades 156, 158 rotate relative to respective sets of compressor vanes 160, 162 to compress or pressurize gas entering through the inlet 132. Referring now only to the HP compressor 118, a single compressor stage 154 includes a plurality of compressor blades 158 (or blades and disks integrated together, referred to as blisks) disposed on a rotor disk 161. The compressor blades extend radially outward from a blade platform to a blade tip relative to the longitudinal centerline axis 112. The compressor vanes 162 are positioned upstream / downstream and adjacent to the rotating compressor blades 158. A rotor disk 161 for a stage of compressor blades 158 is mounted to the high speed shaft 148. A stage of the HPC refers to a single disk of both rotor blades or rotor blades and adjacent stator vanes, both meanings of which may be applied in the context of the present disclosure without loss of clarity.

[0063] The HP turbine 134 has one or two turbine stages 164. In a single turbine stage 164, turbine blades 168 are disposed on a rotor disk 171. The turbine blades extend radially outward from the blade platform to the blade tip relative to the longitudinal centerline axis 112. The HP turbine 134 may also include stator vanes 172. The HP turbine 134 may have an upstream nozzle adjacent to the combustor outlet and an outlet nozzle behind the rotor, or a nozzle upstream of the rotor blades or downstream of the rotor blades.

[0064] Air exiting the HP turbine 134 enters the LP turbine 120, which has multiple stages of rotating blades 170. The LP turbine 120 may have three, four, five, or six stages. In a single LP turbine stage 166 (comprising multiple blades coupled to the LP shaft 122), the turbine blades are disposed on a rotor disk (connected to the LP shaft 122) and extend radially outward from the blade platform to the blade tip relative to the longitudinal centerline axis 112. The LP turbine 120 may also include stator vanes 174. The LP turbine 120 may have an upstream nozzle and an outlet nozzle aft of the stage, followed by an exhaust nozzle 138.

[0065] Figure 2 The turbine engine 100 is Figure 1The engine 10 of the embodiment of the present invention is operated in a similar manner. The airflow leaving the fan section 114 is split so that the first part A1 of the airflow is directed into the inlet 132 of the core passage 115. The core passage 115 includes an LP compressor 121, an HP compressor 118, an LP turbine 120, an HP turbine 134, an LP turbine 120 and an exhaust nozzle 138. The airflow A1 flows to the LP compressor 121, which then supplies the pressurized airflow to the HP compressor 118, which further pressurizes the air. The pressurized airflow from the HP compressor 118 is mixed with fuel and ignited in the combustor 130, thereby generating combustion gases. The HP turbine 134 extracts some work from the combustion gases, which drives the HP compressor 118 to produce self-sustaining combustion. The combustion gases exhausted from the HP turbine 134 enter the LP turbine 120, which extracts additional work to drive the LP compressor 121 and the fan 102 (directly or through the gearbox assembly 150). Exhaust gases from LP turbine 120 exit through exhaust nozzle 138 .

[0066] A second portion A2 of the airflow bypasses the engine core 144 and flows through the bypass passage 113 formed by the fan case 140. The airflow A2 is used to cool portions of the turbine engine 100, particularly the hot portion, and / or to cool or power other aspects of the aircraft. In the context of the turbine engine 100, the hot portion refers to various portions of the turbine engine 100 downstream of the combustion section 128 (e.g., the turbine section 133). Other sources of cooling fluid include, but are not limited to, fluid exhausted from the LP compressor 121 or the HP compressor 118. Although in the Figure 2 Although not shown, engine 100 may include an electric machine operatively coupled to low-speed shaft 122. The electric machine may be coupled to Figure 1 The motor 90 is the same or similar.

[0067] References Figure 1 and Figure 2 The depicted and described gas turbine engines 10 and 100 (also referred to herein as engines 10 and 100) are by way of example only. In other embodiments, the gas turbine engine may have any other suitable configuration, including, for example, any other suitable number or configuration of shafts or spools, fan blades, turbines, compressors, or combinations thereof. The gearbox assembly may have any suitable configuration, including, for example, a star gear configuration, a planetary gear configuration, a single stage, a multi-stage, an epicyclic, a non-episodic, etc., as further described below. The gearbox may have a gear ratio in the range of 3:1 to 4:1, 3:5 to 4:1, 3.25:1 to 3.5:1, or 4:1 to 5:1. The fan assembly may be any suitable fixed pitch assembly or a variable pitch assembly. The gas turbine engine may include Figure 1The fan assembly may be configured in any other suitable manner (e.g., as a fixed pitch fan) and may also be supported using any other suitable fan frame configuration. Aspects of the present disclosure may be incorporated into any other suitable turbine engine (including, but not limited to, turbofan engines, propfan engines, turbojets, turboprop engines, and turboshaft engines).

[0068] Figures 3A to 3C An exemplary sensor assembly for monitoring dynamic torque and torsional frequency in turbine engine 100 is shown. Figures 3A to 3C The sensor assembly monitors the dynamic pressure of the engine flow path (e.g., bypass passage 113 and core passage 115), which is indicative of dynamic torque and torsional frequency in turbine engine 100. The following description includes a return reference to turbine engine 100, and therefore, like numbers represent like features of turbine engine 100. Although shown in turbine engine 100, Figures 3A to 3C The sensor assembly may be provided in the turbine engine 10( Figure 1 ) or other turbine engines.

[0069] Figure 3A An exemplary sensor assembly 200a is shown. The sensor assembly 200a includes a first dynamic pressure sensor 202, a second dynamic pressure sensor 204, and a third dynamic pressure sensor 206. Although three dynamic pressure sensors are depicted and described, more or fewer dynamic pressure sensors may be provided. The first dynamic pressure sensor 202 may be located in the bypass passage 113 between the fan 102 and the outlet guide vane 182. The second dynamic pressure sensor 204 may be located in the core passage 115 between the low pressure compressor and the frame 117. The third dynamic pressure sensor 206 may be located in the core passage 115 between the low pressure turbine 120 and the exhaust nozzle 138.

[0070] Figure 3B An exemplary sensor assembly 200b for monitoring dynamic torque and torsional frequency in a turbine engine 100 is shown. Sensor assembly 200b may be the same or similar to sensor assembly 200a, but with the addition of a first strain gauge 208 and a second strain gauge 210. Although two strain gauges are depicted and described, more or fewer strain gauges may be provided. Figure 3AAs with the aspects of FIG. 1 , more or fewer pressure sensors may be provided. A first strain gauge 208 may be located on a bearing housing of a front bearing assembly 119, also referred to herein as a fan shaft bearing assembly 119. A second strain gauge 210 may be located on a static portion of a gearbox bearing assembly 123 that couples the gearbox assembly 150 to the front frame 125. The first strain gauge 208 may be located on a bearing housing of the front bearing assembly 119 to cancel noise signals from the dynamic pressure sensors 202, 204, 206. The second strain gauge 210 may be located on a stator of the gearbox assembly 150 to monitor the vibration frequency of the gearbox assembly 150.

[0071] Figure 3C An exemplary sensor assembly 200c for monitoring dynamic torque and torsional frequency in a turbine engine 100 is shown. The sensor assembly 200c may be the same or similar to the sensor assembly 200a, but with the addition of a first strain gauge 208 and an accelerometer 212. Figure 3A The first strain gauge 208 may be located on the bearing housing of the front bearing assembly 119, as described with respect to Figure 3B The accelerometer 212 may be located on a frame surrounding the exit guide vane 182. The data or signal from the accelerometer 212 may be used to cancel the noise signals from the dynamic pressure sensors 202, 204, and 206.

[0072] Figures 3A to 3C The sensor assembly 200a, 200b and 200c can provide a dynamic pressure measurement device in the engine flow path (e.g., in the bypass channel 113 and the core channel 115). In some examples, the sensor assembly includes only a single dynamic pressure sensor. The dynamic pressure sensors 202, 204 and 206 are located near or adjacent to the anti-node position on the engine's drive train (e.g., the fan 102, the low-pressure compressor 121 and the low-pressure turbine 120). The dynamic pressure sensors 202, 204 and 206 can detect cavity resonance in the engine flow path. Correlation between regions of the flow path can be performed to take into account various phenomena, such as variable stator vane (VSV) fluctuations. Therefore, the dynamic pressure sensor can monitor the pressure in the engine flow path, which can indicate the dynamic torque in the engine. The signal from the strain gauge or accelerometer can be used to eliminate the noise generated by the engine from the signal.

[0073] Figures 4A to 4D An exemplary sensor assembly for monitoring dynamic torque and torsional frequency in turbine engine 100 is shown. Figures 4A to 4DThe sensor assembly monitors motor parameters indicative of torque and torsional frequency in the engine. The following description includes reference back to turbine engine 100, and thus, like numbers represent like features of turbine engine 100. Although shown in turbine engine 100, Figures 4A to 4D The sensor assembly may be disposed in turbine engine 10 or other turbine engines.

[0074] Figure 4A An exemplary sensor assembly 300a for monitoring dynamic torque and torsional frequency in a turbine engine 100 is shown. The sensor assembly 300a includes an ammeter 302a or other current measuring device and a voltage sensor 304a. Although a single ammeter and a single voltage sensor are shown, multiple ammeters, multiple voltage sensors, or multiple ammeters and multiple voltage sensors may be provided. The ammeter 302a and the voltage sensor 304a are coupled to the motor 190 of the engine 100. The motor 190 may be coupled to the motor 190. Figure 1 The motor 190 may be mounted in the same or similar position, such as on the low speed shaft 122 ( Figure 2 ).

[0075] Figure 4B and 4C An exemplary sensor assembly 300 b for monitoring dynamic torque and torsional frequency in turbine engine 100 is shown. Figure 4C Shows Figure 4B Detailed view 350 of engine 100 of FIG. 100b includes an ammeter 302b or other current measuring device and a voltage sensor 304b. Although a single ammeter and a single voltage sensor are shown, multiple ammeters, multiple voltage sensors, or multiple ammeters and multiple voltage sensors may be provided. Ammeter 302b and voltage sensor 304b are coupled to motor 190a of engine 100. Motor 190a may be coupled to Figure 1 The motor 190a may be the same or similar to the motor 90 of FIG.

[0076] Figure 4D An alternative detail view 350a that may be used in turbine engine 100 is shown with an exemplary sensor assembly 300c for monitoring dynamic torque and torsional frequency in turbine engine 100. Sensor assembly 300c includes an ammeter 302c or other current measuring device and a voltage sensor 304c. Although a single ammeter and a single voltage sensor are shown, multiple ammeters, multiple voltage sensors, or multiple ammeters and multiple voltage sensors may be provided. Ammeter 302c and voltage sensor 304c are coupled to motor 190b of engine 100. Motor 190b may be coupled to Figure 1The motor 190b may be the same or similar to the motor 90 of FIG. 190b. The motor 190b may be mounted to a coupling between the gearbox assembly 150 and the low speed shaft 122.

[0077] Figures 4A to 4D The sensor assemblies 300a, 300b, and 300c can monitor motor parameters indicative of torque and torsional frequency in the engine 100. The voltage sensor provides voltage time domain data and / or current time domain data indicative of torsional vibration in the system. The ammeter provides a signal for eliminating noise (if any) from the voltage sensor signal and indicating the noise of the motor.

[0078] Figure 5 The sensor assembly 200a, 200b, 200c, 300a, 300b or 300c ( Figures 3A-4D ) to calculate the engine 100 ( Figure 2 ) in an exemplary method for determining the dynamic torque in Figure 5 In method 400, at step 402, time domain data is collected. In an example including a pressure sensor (e.g., Figures 3A to 3C ), the data is time domain pressure data, and in examples including a voltage sensor, the data is time domain voltage data (e.g., Figures 4A to 4D ). At step 404, the data is filtered using a bandpass filter. The bandpass filter can be based on several phases. At step 406, the portion of the data that has been filtered through the bandpass filter is plotted as a function of time. At step 408, the data is converted to rad / sec. In an example including a pressure sensor (e.g., Figures 3A to 3C ), the method proceeds to step 410 to calculate the vibration frequency by performing a fast Fourier transform on the data, which allows the dynamic torque to be calculated at step 412. The method may be executed by a processor.

[0079] Figure 5 The method allows the torque to be related to the modal deflection at the sensor location. Figure 5 The outputs of the method may include the magnitude of relative distortion and the frequency of dynamic torque and may allow monitoring limits in the engine to be set based on various response pick-up locations.

[0080] Figures 6A to 6E An exemplary sensor assembly for monitoring dynamic torque and torsional frequency in turbine engine 100 is shown. Figures 6A to 6E The sensor assembly monitors the time lag between sensors to identify torsional displacements indicative of dynamic torque and torsional frequency. The following description includes a return reference to turbine engine 100, and thus, like numbers represent like features of turbine engine 100. Although shown in turbine engine 100, Figures 6A to 6EThe sensor assembly may be disposed in turbine engine 10 or other turbine engines.

[0081] Fig. 6A An exemplary sensor assembly 600a is shown. The sensor assembly 600a includes a first magnetic sensor 602, a second magnetic sensor 604, a third magnetic sensor 606, and a fourth magnetic sensor 608. Although four magnetic sensors are depicted and described, more or fewer magnetic sensors may be provided. However, in order to determine the time lag, at least two magnetic sensors are required, as described in more detail below. The first magnetic sensor 602 is located at the front bearing assembly 119 on the fan shaft of the fan 102. The second magnetic sensor 604 is located between the front bearing assembly 111 in front of the low-pressure compressor 121 and the rear bearing assembly 127 behind the low-pressure compressor 121. The third magnetic sensor 606 is located between the first bearing assembly 129 in front of the high-pressure turbine 134 and the second bearing assembly 135 behind the high-pressure turbine 134. The fourth magnetic sensor 608 is located at the motor 190. The magnetic sensors 602, 604, 606, 608 are connected to a processor 611, which is configured to process real-time data collected by the magnetic sensors.

[0082] Each of the magnetic sensors 602, 604, 606, 608 is relative to the longitudinal centerline axis 112 ( Figure 2 ) are located at substantially the same radial position, but are axially spaced apart along the length of the fan shaft 157 and the low-speed shaft 122. Magnetic material or a magnetic indicator is placed on a rotating shaft (e.g., the fan shaft 157 or the low-speed shaft 122) at each sensor location. When the magnetic material passes by a corresponding sensor (e.g., by rotation of the rotating shaft), the sensor generates a signal or pulse indicating one rotation of the rotating shaft. The processor 611 can process the signal and identify a time lag between the axially spaced magnetic sensors 602, 604, 606, 608, which indicates a rotor torsion pattern. To identify the time lag, the signal is compared to a predefined reference time based on rotor speed or similar data.

[0083] Figure 6B and Figure 6C An exemplary sensor assembly 600b is shown. The sensor assembly 600b includes a first magnetic sensor 610 and a second magnetic sensor 612. The first magnetic sensor 610 is located at an outer diameter D of the motor 190. EM The outer diameter D of the motor 190 EM A plurality of first magnetic protrusions 614 may be included. The second magnetic sensor 612 is located at an outer diameter D of the gearbox assembly 150. GB The outer diameter D of the gearbox assembly 150 is GBA plurality of second magnetic protrusions 616 may be included. The first magnetic sensor 610 and the second magnetic sensor 612 are coupled to a processor 618 that is configured to process data acquired by the magnetic sensors. Although two magnetic sensors are shown and described, more or fewer magnetic sensors may be provided.

[0084] In the sensor assembly 600b, the magnetic sensors 610, 612 may be magnetic induction sensors. The sensor assembly 600b monitors and calculates the most extreme tangential displacement that may be produced on the gearbox assembly 150 and the motor 190. The magnetic protrusions 614, 616 may produce depressions and protrusions (e.g., similar to gear teeth) on the gearbox housing (e.g., a plurality of second magnetic protrusions 616) and the generator housing (e.g., a plurality of first magnetic protrusions 614). The first magnetic sensor 610 detects the differential magnetic flux of the channel of each depression and protrusion on the gearbox housing, and the second magnetic sensor 612 detects the differential magnetic flux of the channel of each depression and protrusion on the motor housing. The relative sensor deflection produced by the depression and protrusion of the protrusion identifies the amplitude exceeding the nominal tangential deflection of the gearbox assembly 150 and the motor 190, which indicates torsional vibration. Processor 618 calculates the relative clocking of gearbox assembly 150 and motor 190 over predefined duration samples to identify acceptable torsional amplitudes present in engine 100. Processor 618 identifies sudden increases in amplitude variations over the duration samples to indicate the presence of torsional vibrations.

[0085] Fig.6D and Fig. 6E An exemplary sensor assembly 600c is shown. The sensor assembly 600c includes a first light detector 620 and a second light detector 622. The first light detector 620 is located at the first bearing assembly 139 of the gearbox assembly 150. The first bearing assembly 139 includes a plurality of bearings, such as a first bearing 139a and a second bearing 139b. The second light detector 622 is located at the second bearing assembly 135 in front of the motor 190. The second bearing assembly 135 includes a plurality of bearings, such as a first bearing 135a and a second bearing 135b. The first light detector 620 and the second light detector 622 are coupled to a processor 624, which is configured to process data collected by the detectors. Although two light detectors are shown and described, more or fewer light detectors may be provided.

[0086] The sensor assembly 600c utilizes the principle of blade tip timing to produce a system indicating torsional vibration. The sensor assembly 600c measures the passing events of each bearing of the bearing assemblies 139 and 135 at a specific angular position. The sensor assembly 600c includes a light source or laser source, a photodetector / amplifier, and a light detector (e.g., a first light detector 620 and a second light detector 622). In some examples, the light detector can be mounted to illuminate a roller bearing or a ball bearing (e.g., a bearing of the bearing assemblies 139 and 135) and capture light reflections when the bearing passes near the light detector. In some examples, the light detector can be a pitch capture detector that is mounted in a manner that depends on the bearing to chop the light signal when the bearing passes between the pitch capture detectors. The differential measurement of the bearing passing rate between the ball bearing at the front bearing assembly 139 and the roller bearing at the rear bearing assembly 135 provides an indication of the torsion in the low-speed shaft 122.

[0087] Figures 6A to 6C The sensor assembly includes a magnetic pickup to detect the time lag from one sensor position to another sensor position. That is, at each sensor position, a magnetic material (e.g., a magnetic protrusion, strip, etc.) is detected by the sensor. Fig.6D and Fig. 6E The sensor assembly 600c includes a light detector that detects the time lag from one sensor position to another sensor position. Each sensor of the above embodiment is placed at the same radial position. In the absence of torsional displacement, the time for the corresponding sensors to pass through the magnetic material or generate the light signal should be the same on the rotating shaft (e.g., the low speed shaft 122 and / or the fan shaft 157). FIG. 6A to FIG. 6E The sensors of the sensor assembly detect time lags between sensors placed along the axial length of the engine. The time lags indicate torsional displacement, which indicates dynamic torque and torsional frequency of the engine 100.

[0088] Figure 7 An exemplary sensor assembly 700 for monitoring dynamic torque and torsional frequency in a turbine engine 100 is shown. The following description includes references back to the turbine engine 100, and thus, like numbers refer to like features of the turbine engine 100. Although in the turbine engine 100 ( Figure 2 ), but Figure 7 The sensor assembly may be provided in the turbine engine 10( Figure 1) or other turbine engines. Sensor assembly 700 includes a first incremental rotary encoder 702 and a second incremental rotary encoder 704. Although two incremental rotary encoders are depicted and described, more or fewer incremental rotary encoders may be provided. First incremental rotary encoder 702 is located on fan shaft 157 that couples fan 102 to low-speed shaft 122. Second incremental rotary encoder 704 is located on low-speed shaft 122.

[0089] The sensor assembly 700 requires at least one incremental rotary encoder (e.g., second incremental rotary encoder 704) mounted on the low-speed shaft. Angular displacement data in the time domain is collected by the incremental rotary encoder to indicate torsional vibration of the system. Additional incremental rotary encoders can be mounted on the input shaft of the gearbox assembly (e.g., first incremental rotary encoder 702) to understand transmission errors (e.g., changes in angular displacement between the input shaft of the gearbox assembly 150 (e.g., low-speed shaft 122) and the output shaft of the gearbox assembly 150 (e.g., fan shaft 157)), which is the cause of torsional vibration in the system.

[0090] Figure 8 A method 800 for monitoring dynamic torque and torsional frequency is shown, comprising: Figure 7 The sensor assembly 700 or Figures 6A to 6E At step 802, the method includes setting a sensor assembly 600a, 600b, 600c on the engine 100. Figure 7 The incremental rotary encoder described (or magnetic sensor or optical sensor, as described in Figures 6A to 6E ), and at step 804, the system is modeled. At step 806, encoder time domain data is collected using the sensor assembly 700. At step 808, the data is filtered, and at step 810, the data is transformed to enhance visualization, and at step 812, the method 800 performs a fast Fourier transform on the data, which allows for visualization, plotting, and calculation of dynamic torque at steps 814 and 816. The method may be executed by a processor. Figure 8 The method can be adopted on a model engine or a virtual model to indicate the dynamic torque in the engine. Multiple sensors can be introduced on the stationary and rotating parts in the virtual model. The virtual model can be used to simulate the angular displacement at all response frequencies and resonant frequencies, and therefore, the torsional vibration at any frequency can be calculated.

[0091] Figures 9 to 11B A method for monitoring a turbine engine 100 ( Figure 2) in an exemplary sensor assembly for dynamic torque and torsional frequency in a turbine engine 100. The following description includes reference back to turbine engine 100, and thus, like numbers represent like features of turbine engine 100. Although shown in turbine engine 100, Figures 9 to 11B The sensor assembly may be provided in the turbine engine 10( Figure 1 ) or other turbine engines.

[0092] Fig. 9 An exemplary sensor assembly 900a is shown. The sensor assembly 900a includes an optical sensor 902 mounted at a target or strategic location. For example, Fig. 9 The target location shown is on the most downstream stage of the low pressure turbine 120. Other strategic or target locations are also contemplated. Although a single optical sensor 902 is shown, more optical sensors may be provided. In one example, the optical sensor 902 may be located on a stator frame downstream of the last stage of the low pressure turbine 120. In another example, the optical sensor 902 may be located on a housing extending radially inward from the frame so that the rotating bolts of the low pressure turbine are in view (e.g., in Fig.12 In another example, the optical sensor 902 may be located upstream of the first stage of the low pressure turbine 120 on a frame in front of the low pressure turbine 120 (referred to as the turbine center frame). In such a location, the optical sensor 902 may pick up Fig.12 In another example, the optical sensor 902 may be in a front section of the engine, such as near a supercharger or compressor stage, and mounted to the frame 117 ( Figure 3A ) to measure a rotating bolt similar to the rotating bolt discussed above with respect to the low pressure turbine 120. In some examples, the optical sensor 902 may be located on the stator of the gearbox assembly 150.

[0093] Fig.10 A method 1000 using the sensor assembly 900a is shown. At step 1002, an optical sensor is placed at a target location. At step 1004, torsional fluctuations are calculated based on data from the optical sensor. At step 1006, the motor 190 ( Figure 2 ). At step 1008, a correlation diagram is generated between the torsional fluctuations (e.g., data from step 1004) and the current fluctuations (e.g., data from step 1006). Step 1010 allows monitoring of the rotor (e.g., Figure 2 torsion in the low speed shaft 122).

[0094] In the sensor assembly 900a ( Fig. 9 ), the optical sensor 902 ( Fig. 9) detects a belt or patch on a rotating component disposed at a target location. Based on the rotor speed and the moment of the blip caused by the passage of the belt or patch, the torsional fluctuations can be calculated. The sensor assembly 900a can also record the current fluctuations in the motor, and a correlation diagram can be generated between the torsion in the rotor (as sensed by the optical sensor 902) and the current fluctuations in the motor. The correlation diagram can then be used to monitor the torque in the rotor of the live engine.

[0095] Fig.11A and Fig. 11B An exemplary sensor assembly 900b is shown. The sensor assembly 900b can be combined with the sensor assembly 900a ( Fig. 9 ) is the same or similar to the sensor assembly 900b, but with the addition of a second optical sensor. That is, the sensor assembly 900b includes a first optical sensor 904 and a second optical sensor 906. The optical sensors 904 and 906 are spaced apart on the low-speed shaft 122. For example, the first optical sensor 904 can be located at the front end of the low-speed shaft 122, and the second optical sensor 906 can be located at the rear end of the low-speed shaft 122. The first optical sensor 904 and the second optical sensor 906 can be located at the same radial position relative to the low-speed shaft 122. Fig.11A and Fig. 11B shows how optical sensors allow the monitoring of torsion in a low speed shaft. Fig.11A In the embodiment, when there is no torsion in the low speed shaft, both the first optical sensor 904 and the second optical sensor 906 will monitor the position on the longitudinal axis 910 of the low speed shaft 122. If there is torsion, such as Fig. 11B As shown, the twist will cause the first optical sensor 904 and the second optical sensor 906 to monitor positions at different longitudinal lines of the low speed shaft 122. For example, the second optical sensor 906 may continue to monitor the position on the longitudinal axis 910, while the first optical sensor 904 may monitor a position radially offset from the position on another longitudinal axis 912. This is due to the twist in the low speed shaft 122. The angle 914 between the longitudinal axes 910 and 912 may be related to the twist in the low speed shaft 122.

[0096] Figures 12 to 14B A method for monitoring a turbine engine 100 ( Figure 2 ) in an exemplary sensor assembly for dynamic torque and torsional frequency in a turbine engine 100. The following description includes reference back to turbine engine 100, and thus, like numbers represent like features of turbine engine 100. Although shown in turbine engine 100, Figures 12 to 14B The sensor assembly may be provided in the turbine engine 10( Figure 1 ) or other turbine engines.

[0097] Fig.12An exemplary sensor assembly 1200a is shown. The sensor assembly 1200a includes a high-speed camera 1202 mounted at a target or strategic location. For example, Fig.12 The target location shown is on the most downstream stage of the low pressure turbine 120. Other strategic or target locations are also contemplated. For example, the high speed camera 1202 may be located on a frame downstream of the low pressure turbine 120, on a frame upstream of the low pressure turbine 120, or about the optical sensor 902 ( Fig. 9 ) described above. As with the optical sensor 902, the high-speed camera 1202 must be mounted to a static (e.g., non-rotating) component of the engine. Although only one high-speed camera 1202 is shown, more high-speed cameras may be provided. A screen 1204 may be included in front of the high-speed camera 1202 to protect the high-speed camera 1202. Fig.13 A method 1300 using the sensor assembly 1200a is shown. At step 1302, a high-speed camera acquires an image. At step 1304, the image is processed to generate an indication of torsional fluctuations. At step 1306, the motor 190 ( Figure 2 ). At step 1308, a correlation diagram is generated between the torsional fluctuations (e.g., data from step 1304) and the current fluctuations (e.g., data from step 1306). Step 1310 allows monitoring of the rotor (e.g., Figure 2 torsion in the low speed shaft 122).

[0098] In sensor assembly 1200a, a high-speed camera 1202 detects a color-coded image on a rotating component disposed at a target location. Based on the rotor speed and the moment of the light spot caused by the passage of the belt or patch, the torsional fluctuations can be calculated. Sensor assembly 1200a can also record the current fluctuations in the motor, and a correlation diagram can be generated between the torsion in the rotor (as sensed by high-speed camera 1202) and the current fluctuations in the motor. The correlation diagram can then be used to monitor the torque in the rotor of a live engine. High-speed camera 1202 allows high-speed camera imaging combined with image processing for torsional measurement.

[0099] Fig.14A and 14B An exemplary sensor assembly 1200b is shown. Sensor assembly 1200b can be the same or similar to sensor assembly 1200a, except that a second high-speed camera is added. That is, sensor assembly 1200b includes a first high-speed camera 1206 and a second high-speed camera 1208. The optical sensors are spaced apart on the low-speed shaft 122. For example, the first high-speed camera 1206 can be located at the front end of the low-speed shaft 122, and the second high-speed camera 1208 can be located at the rear end of the low-speed shaft 122. The first high-speed camera 1206 and the second high-speed camera 1208 can be located at the same radial position relative to the low-speed shaft 122. Fig.14A and Fig. 14B shows how a high-speed camera allows the monitoring of torsion in a low-speed shaft. Fig.14A In the embodiment, when there is no torsion in the low speed shaft, both the first high speed camera 1206 and the second high speed camera 1208 will monitor the position on the longitudinal axis 1210 of the low speed shaft 122. If there is torsion, such as Fig. 14B As shown, the twist will cause first high-speed camera 1206 and second high-speed camera 1208 to monitor positions at different longitudinal lines of low speed shaft 122. For example, first high-speed camera 1206 may continue to monitor positions on longitudinal axis 1210, while second high-speed camera 1208 may monitor positions radially offset from positions on another longitudinal axis 1212. This is due to the twist in low speed shaft 122. Angle 1214 between longitudinal axes 1210 and 1212 may be related to the twist in low speed shaft 122.

[0100] Any of some or all of the above-described sensor components may be used alone, or in combination with some or all of the other sensor components described herein.

[0101] Thus, each of the above sensor assemblies provides a system for monitoring and detecting parameters indicative of torsional vibrations in the engine. Once torsional vibrations in the low-speed shaft are detected, actions can be taken to damp the vibrations. These actions can provide mechanical damping (e.g., a clutch or torsional damper) or electrical damping (e.g., a varying electrical load). Damping can reduce vibrations in the low-speed shaft 122 or can eliminate vibrations in the low-speed shaft 122.

[0102] exist Fig.15 Several exemplary systems that perform this action are shown in . Fig.15 Shows Figure 2 A portion of an engine with features removed for clarity. Figure 2 The engine is shown, but Fig.15 The system can be Figure 1 In addition, it should be understood that Fig.15 The system for taking action described may be found in FIG. 3A to FIG. 14B Any sensor system described may be employed.

[0103] refer to Fig.15 , engine 1500 includes a low speed shaft 122 and an electric machine 90. A first damping system 1502 that damps vibrations in the low speed shaft 122 can be associated with the electric machine 90. The first damping system 1502 can vary the electrical load demand on the electric machine 90 to provide damping of torsional vibrations in the low speed shaft 122.

[0104] Continue to refer Fig.15, the second damping system 1504 can be a variable torsional damper located on the low-speed shaft 122, longitudinally close to the motor 90. The torsional damper can be any known torsional damper. The torsional damper provides damping of torsional vibrations in the low-speed shaft 122. The torsional damper can provide variable damping that can be controlled by a feedback loop, as described below. In other examples, the torsional damper can be a fixed damping value. In such an example, since the torsional damping applied to the low-speed shaft 122 will be fixed, a feedback loop may not be employed. The torsional damper can be a mechanical damper.

[0105] The third damping system 1506 may be a clutch on the rear end of the low speed shaft 122. The clutch of the third damping system 1506 may be longitudinally located in the bearing (e.g., about Fig. 6A The fourth damping system 1508 may be a clutch on the front end of the low speed shaft 122. The clutch of the fourth damping system 1508 may be located longitudinally of the gearbox assembly (e.g., about Fig. 6A gearbox assembly 150 described above) and a low pressure compressor (e.g., Fig. 6A 121). Either or both clutches may be provided with the low speed shaft 122. The clutches provide damping of torsional vibrations in the low speed shaft 122. Other locations of the clutches along the low speed shaft 122 are contemplated.

[0106] The damping system described above can be used in a feedback loop with the monitoring system of the present disclosure. That is, the monitoring of the torsional vibration can be continuous, and the torsional vibration level can be provided to the damping system. If the vibration level is above a predetermined value, the damping system can be activated to damp the vibration in the low speed shaft. The monitoring continues and the updated vibration level is detected. The damping system can continue to be used to maintain the vibration level within an acceptable range. Thus, a closed feedback loop is provided for monitoring and damping the vibration in the low speed shaft.

[0107] Any of the above damping systems for implementing remedial action may be provided alone or in combination with other damping systems. Additionally, when the engine stops functioning, a maintenance message may be generated to alert the user that further remediation is required. The maintenance message may indicate the immediacy of maintenance required based on the measured torsional vibrations.

[0108] Thus, the sensor assembly of the present disclosure provides a system for monitoring and detecting parameters that may be indicative of torsional dynamics (i.e., torsional vibration performance in an engine). The sensor assembly monitors various parameters, such as dynamic pressure in the engine flow path, motor parameters, torsional displacement via time lag between magnetic sensors or photodetectors, optical or high-speed camera images, or torsional displacement via incremental rotary encoders. Monitoring various parameters in a model, whether physical or virtual, allows monitoring of torsional vibrations in a live engine and may allow corrective or remedial action to be taken if the torsional vibrations are outside a predetermined acceptable range.

[0109] Further aspects are provided by the subject matter of the following clauses.

[0110] A gas turbine engine having a sensor assembly configured to detect torsional vibrations, the gas turbine engine comprising: a fan section having a plurality of fan blades, the fan section configured to generate an airflow through the gas turbine engine; an airflow passage having a core passage and a bypass passage, the core passage being configured to guide a first portion of the airflow from an inlet to an exhaust nozzle, the bypass passage being separated from the core passage and configured to guide a second portion of the airflow around the core passage; a low-speed shaft coupled to the plurality of fan blades and configured configured to rotate the plurality of fan blades; and a sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising a plurality of dynamic pressure sensors in the airflow passage, the plurality of dynamic pressure sensors detecting dynamic pressures of the airflow passage, the dynamic pressures indicating the torsional vibrations in the low-pressure shaft, the sensor assembly being configured to monitor torsional vibrations in the gas turbine engine; and a damping system configured to damp the torsional vibrations in the low-speed shaft based on a correlation between measured torsional vibrations and experimental torsional vibrations.

[0111] According to the gas turbine engine of the preceding clause, the plurality of dynamic pressure sensors includes at least one dynamic pressure sensor in the core passage and at least one dynamic pressure sensor in the bypass passage.

[0112] A gas turbine engine according to any preceding clause, further comprising a low pressure compressor and a low pressure turbine.

[0113] The gas turbine engine according to the preceding clause, wherein at least one dynamic pressure sensor of said plurality of dynamic pressure sensors is located at said fan section, said low pressure compressor or said low pressure turbine.

[0114] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0115] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0116] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0117] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0118] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0119] A gas turbine engine as described in any preceding clause, wherein the plurality of dynamic pressure sensors comprises a first dynamic pressure sensor, a second dynamic pressure sensor, and a third dynamic pressure sensor.

[0120] A gas turbine engine as claimed in any preceding clause, wherein the first dynamic pressure sensor is located in the bypass passage.

[0121] A gas turbine engine as described in any preceding clause, wherein the first dynamic pressure sensor is located aft of the fan section.

[0122] A gas turbine engine according to any preceding clause, further comprising an outlet guide vane, said first dynamic pressure sensor being located forward of said outlet guide vane.

[0123] A gas turbine engine as described in any preceding clause, wherein the second dynamic pressure sensor is located in the core passage.

[0124] The gas turbine engine of any preceding clause, further comprising a low pressure compressor, the second dynamic pressure sensor being located after the low pressure compressor.

[0125] A gas turbine engine as claimed in any preceding clause, wherein said third dynamic pressure sensor is located in said core passage.

[0126] The gas turbine engine according to any preceding clause, further comprising a low pressure turbine, said third dynamic pressure sensor being located after said low pressure turbine and in front of said exhaust nozzle.

[0127] The gas turbine engine of any preceding clause, the sensor assembly further comprising at least one strain gauge configured as a filter to filter noise from the plurality of dynamic pressure sensors.

[0128] The gas turbine engine of any preceding clause, further comprising a bearing assembly having a bearing housing support, the bearing assembly coupled to the fan section, the at least one strain gauge being located on the bearing housing support of the bearing assembly.

[0129] A gas turbine engine according to any preceding clause, further comprising a forward bearing assembly located forward of said low pressure compressor, said at least one strain gauge being located on a bearing housing of said forward bearing assembly.

[0130] The gas turbine engine of any preceding clause, further comprising a gearbox assembly coupled between the fan section and the low speed shaft, the at least one strain gauge comprising a second strain gauge located on a stator of the gearbox assembly.

[0131] The gas turbine engine of any preceding clause, the sensor assembly further comprising an accelerometer configured as a filter to filter noise from the plurality of dynamic pressure sensors.

[0132] A gas turbine engine according to any preceding clause, further comprising an exit guide vane having a frame, said accelerometer being located on said frame of said exit guide vane.

[0133] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0134] A turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0135] A gas turbine engine, comprising: a fan section having a plurality of fan blades, the fan section being configured to generate an airflow through the gas turbine engine; an airflow passage having a core passage and a bypass passage, the core passage being configured to direct a first portion of the airflow from an inlet to an exhaust nozzle, the bypass passage being separate from the core passage and configured to direct a second portion of the airflow around the core passage; a low-speed shaft coupled to the plurality of fan blades and configured to rotate the plurality of fan blades; a sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising at least one dynamic pressure sensor in the airflow passage and configured to detect a dynamic pressure of the airflow passage, the dynamic pressure being indicative of the torsional vibrations in the low-speed shaft; and a strain gauge or accelerometer coupled to a static component of the gas turbine engine, the strain gauge or accelerometer being configured as a filter to filter noise from the at least one dynamic pressure sensor.

[0136] The gas turbine engine of any preceding clause, further comprising the strain gauge and a bearing assembly having a bearing housing support, the bearing assembly coupled to the fan section, the strain gauge being located on the bearing housing support of the bearing assembly.

[0137] A gas turbine engine according to any preceding clause, further comprising a forward bearing assembly located forward of said low pressure compressor, said strain gauge being located on a bearing housing of said forward bearing assembly.

[0138] The gas turbine engine of any preceding clause, further comprising a gearbox assembly coupled between the fan section and the low speed shaft, the strain gauge comprising a second strain gauge located on a stator of the gearbox assembly.

[0139] A gas turbine engine as described in any preceding clause, said sensor assembly further comprising said accelerometer.

[0140] A gas turbine engine according to any preceding clause, further comprising an exit guide vane having a frame, said accelerometer being located on said frame of said exit guide vane.

[0141] A gas turbine engine as described in any preceding clause, wherein said at least one dynamic pressure sensor is located in said core passage.

[0142] A gas turbine engine according to any preceding clause, further comprising said strain gauge and said accelerometer.

[0143] A gas turbine engine as described in any preceding clause, wherein the at least one dynamic pressure sensor comprises a first dynamic pressure sensor located in the core passage and a second dynamic pressure sensor located in the bypass passage.

[0144] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0145] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0146] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0147] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0148] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0149] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0150] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0151] A turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0152] A gas turbine engine comprises: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; an electric motor operably coupled to the low-speed shaft; and a sensor assembly coupled to the electric motor and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising a voltage sensor configured to sense time domain data indicative of torsional vibrations of the low-speed shaft.

[0153] The gas turbine engine according to the preceding clause, wherein the time domain data is voltage time domain data or current time domain data.

[0154] A gas turbine engine according to any preceding clause, further comprising an ammeter configured to cancel noise from the electric machine.

[0155] A gas turbine engine as claimed in any preceding clause, wherein the current meter is coupled to the electric machine.

[0156] According to the gas turbine engine of the preceding clause, the electric machine is coupled to the low speed shaft at a rearward end of the low speed shaft.

[0157] A gas turbine engine as described in any preceding clause, further comprising a gearbox assembly coupled between said fan and said low speed shaft, said electric machine being coupled to said gearbox assembly.

[0158] A gas turbine engine according to any preceding clause, further comprising a gearbox assembly coupled between said fan and said low speed shaft, said electric machine being coupled to a coupling region between said gearbox assembly and said low speed shaft.

[0159] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0160] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0161] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0162] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0163] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0164] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0165] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0166] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0167] A gas turbine engine comprises: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; an electric motor operably coupled to the low-speed shaft; and a sensor assembly coupled to the electric motor and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising a plurality of magnetic sensors coupled to the gas turbine engine and configured to detect time lags between the plurality of magnetic sensors, the time lags indicating torsional vibrations of the low-speed shaft.

[0168] The gas turbine engine according to the preceding clause, wherein at least two magnetic sensors of the plurality of magnetic sensors are located at substantially the same radial position relative to a longitudinal centerline axis of the gas turbine engine.

[0169] A gas turbine engine as in any preceding clause, wherein at least two of said plurality of magnetic sensors are axially spaced apart along a longitudinal centerline axis of said gas turbine engine.

[0170] The gas turbine engine according to the preceding clause, further comprising a fan shaft operably coupled between the fan and the low speed shaft, the plurality of magnetic sensors comprising a magnetic sensor coupled to a fan shaft bearing assembly of the fan shaft.

[0171] The gas turbine engine of any preceding clause, further comprising a first aft bearing assembly positioned adjacent the low pressure compressor, the plurality of magnetic sensors comprising a magnetic sensor coupled to the first aft bearing assembly.

[0172] The gas turbine engine of any preceding clause, further comprising a second aft bearing assembly positioned adjacent the high pressure turbine, the plurality of magnetic sensors comprising a magnetic sensor coupled to the second aft bearing assembly.

[0173] A gas turbine engine as described in any preceding clause, wherein said plurality of magnetic sensors comprises a magnetic sensor coupled to said electric machine.

[0174] The gas turbine engine according to any preceding clause, further comprising a magnetic strip coupled to the rotating part and aligned with each magnetic sensor of the plurality of magnetic sensors, each magnetic sensor being configured to generate a signal when the magnetic strip passes by the respective magnetic sensor.

[0175] A gas turbine engine as described in any preceding clause, wherein the plurality of magnetic sensors comprises magnetic materials, magnetic indicators, magnetic protrusions, magnetic inductive sensors, or a combination thereof.

[0176] The gas turbine engine according to the preceding clause, further comprising a filter configured to process said signal.

[0177] A gas turbine engine as claimed in any preceding clause, wherein the magnetic sensor is a magnetic induction sensor.

[0178] The gas turbine engine according to any of the preceding clauses, further comprising a gearbox assembly, the gearbox assembly comprising a plurality of magnetic protrusions extending from an outer surface of the gearbox assembly, a first magnetic sensor of the plurality of magnetic sensors being configured to sense a change in magnetic flux when the plurality of magnetic protrusions pass by the first magnetic sensor.

[0179] The gas turbine engine of any preceding clause, further comprising a plurality of magnetic protrusions extending from an outer surface of the electric machine, a second magnetic sensor of the plurality of magnetic sensors being configured to sense a change in magnetic flux when the plurality of magnetic protrusions pass by the second magnetic sensor.

[0180] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0181] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0182] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0183] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0184] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0185] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0186] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0187] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0188] A method for calculating the dynamic torque of a low-speed shaft in a gas turbine engine, the method comprising: coupling a sensor assembly to the low-speed shaft of the gas turbine engine; collecting time domain data; filtering at least a portion of the time domain data to generate filtered data; plotting the filtered data as a function of time; processing the filtered data to obtain a vibration frequency of the low-speed shaft; and calculating the dynamic torque of the low-speed shaft based on the vibration frequency of the shaft.

[0189] The method of the preceding clause, further comprising filtering the data using a bandpass filter to generate filtered data.

[0190] The method of any preceding clause, wherein collecting time-domain data comprises sensing voltage using one or more voltage sensors, the time-domain data comprising time-domain voltage data.

[0191] Method according to the preceding clause, wherein said one or more voltage sensors are coupled to an electric machine of said gas turbine engine.

[0192] A method as in any preceding clause, further comprising filtering noise from the time domain voltage data using one or more current meters coupled to the motor.

[0193] The method of any preceding clause, wherein collecting time-domain data comprises sensing pressure using one or more pressure sensors, the time-domain data comprising time-domain pressure data.

[0194] A method according to the preceding clause, wherein sensing pressure comprises sensing pressure in a bypass passage of the gas turbine engine between a fan and an outlet guide vane, sensing pressure in a core passage of the gas turbine engine between a low-pressure compressor and a frame, sensing pressure in the core passage of the gas turbine engine between a low-pressure turbine and an exhaust nozzle, or any combination thereof.

[0195] A method as in any preceding clause, further comprising removing noise from the time domain pressure data using data from one or more strain gauges.

[0196] A method as in any preceding clause, further comprising removing noise from the time domain pressure data using data from one or more strain gauges and one or more accelerometers.

[0197] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft based on the calculated dynamic torque.

[0198] A method as in any preceding clause, further comprising mechanically damping the vibrations of the low speed shaft.

[0199] A method as in any preceding clause, further comprising electrically damping said vibrations of said low speed shaft.

[0200] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft using one or more clutches coupled to the low speed shaft.

[0201] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft with a torsional damper coupled to the low speed shaft.

[0202] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft by varying a load of an electric machine coupled to the low speed shaft.

[0203] A method as claimed in any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0204] A method as claimed in any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0205] A gas turbine engine, comprising: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft, the low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; a motor, the motor operably coupled to the low-speed shaft; and a sensor assembly, the sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising a front light detector coupled to the gas turbine engine and a rear light detector coupled to the gas turbine engine, the sensor assembly being configured to detect a difference between a first signal from the front light detector and a second signal from the rear light detector, the difference indicating the torsional vibrations in the low-speed shaft.

[0206] The gas turbine engine according to the preceding clause, further comprising a gearbox assembly coupled between the fan and the low speed shaft, the front light detector coupled to a bearing assembly of the gearbox assembly, and the rear light detector coupled to a bearing assembly of the motor.

[0207] According to the gas turbine engine described in any preceding item, the front light detector and the rear light detector are each configured to detect reflected light from a light source, or are configured to detect the absence of light from a light source when multiple bearings rotate past the front light detector and the rear light detector respectively.

[0208] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0209] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0210] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0211] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0212] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0213] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0214] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0215] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0216] A gas turbine engine, comprising: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft, the low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; a fan shaft, the fan shaft coupled between the fan and the low-speed shaft; a motor, the motor operably coupled to the low-speed shaft; and a sensor assembly, the sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly having a front incremental rotary encoder coupled to the fan shaft and a rear incremental rotary encoder coupled to the low-speed shaft, each of the front incremental rotary encoder and the rear incremental rotary encoder being configured to detect an angular displacement of the low-speed shaft indicative of torsional vibrations.

[0217] The gas turbine engine according to the preceding clause, further comprising a gearbox assembly including another incremental rotary encoder configured to detect a transmission error indicative of said torsional vibration of said low speed shaft.

[0218] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0219] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0220] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0221] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0222] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0223] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0224] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0225] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0226] A method for monitoring dynamic torque and torsional frequency in a low-speed shaft of a gas turbine engine, the method comprising: positioning one or more sensors on the low-speed shaft of the gas turbine engine to generate a sensor assembly on the gas turbine engine; modeling the sensor assembly on a virtual gas turbine engine on a computer having a processor; collecting data using the sensor assembly on the gas turbine engine; filtering the data to generate filtered data; processing the filtered data to determine the frequency and amplitude of torsional vibration of the low-speed shaft; plotting the frequency and amplitude of torsional vibration; and calculating dynamic torque based on the frequency and amplitude of the torsional vibration.

[0227] A method as described in the preceding clause, wherein the one or more sensors are one or more incremental rotary encoders.

[0228] A method as in any preceding clause, wherein the one or more sensors are one or more magnetic sensors.

[0229] A method as in any preceding clause, wherein the one or more sensors are one or more light detectors.

[0230] A method as recited in the preceding clause, further comprising filtering the data using a bandpass filter to generate filtered data.

[0231] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft based on the calculated dynamic torque.

[0232] A method as in any preceding clause, further comprising mechanically damping the vibrations of the low speed shaft.

[0233] A method as in any preceding clause, further comprising electrically damping said vibrations of said low speed shaft.

[0234] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft using one or more clutches coupled to the low speed shaft.

[0235] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft with a torsional damper coupled to the low speed shaft.

[0236] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft by varying a load of an electric machine coupled to the low speed shaft.

[0237] A method as claimed in any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0238] A method as claimed in any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0239] A gas turbine engine comprises: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft, the low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; and a sensor assembly, the sensor assembly being coupled to the gas turbine engine, the sensor assembly having an optical sensor coupled to a static component of the gas turbine engine and an indicator coupled to a rotating component of the gas turbine engine, the optical sensor being configured to detect the indicator as the indicator rotates past the optical sensor to detect a rotor speed of a rotating object, and to detect the passage of the indicator to indicate the presence of torsional fluctuations in the low-speed shaft.

[0240] According to the gas turbine engine of the preceding clause, the rotating component is the most downstream stage of the low-pressure turbine.

[0241] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located on a stator frame downstream of a most downstream stage of the low pressure turbine.

[0242] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located on a casing extending radially inwardly from a frame of the gas turbine engine.

[0243] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located upstream of a first stage of the low pressure turbine on a frame forward of the low pressure turbine.

[0244] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located in a forward section of the gas turbine engine.

[0245] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located adjacent a low pressure compressor of the gas turbine engine.

[0246] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located in the gas turbine engine so as to sense a rotating bolt within the gas turbine engine.

[0247] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor is located on a stator of a gearbox assembly of the gas turbine engine.

[0248] A gas turbine engine as described in any preceding clause, wherein the optical sensor comprises a plurality of optical sensors.

[0249] A gas turbine engine as claimed in any preceding clause, wherein the optical sensor comprises at least two optical sensors located at substantially the same radial position relative to a centreline axis of the gas turbine engine.

[0250] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0251] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0252] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0253] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0254] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0255] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0256] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0257] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0258] A gas turbine engine comprises: a fan; a low-pressure turbine; a low-pressure compressor; a low-speed shaft, the low-speed shaft drivingly connecting the low-pressure turbine to the low-pressure compressor and the fan; and a sensor assembly coupled to the gas turbine engine, the sensor assembly having a high-speed camera coupled to a static component of the gas turbine engine and an indicator coupled to a rotating component of the gas turbine engine, the high-speed camera being configured to detect the indicator as the indicator rotates past the high-speed camera, and being configured to indicate the presence of torsional fluctuations in the low-speed shaft.

[0259] The gas turbine engine according to the preceding clause, further comprising a screen in front of the high-speed camera for protecting the high-speed camera.

[0260] According to the gas turbine engine of the preceding clause, the rotating component is the most downstream stage of the low-pressure turbine.

[0261] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located on a stator frame downstream of a most downstream stage of the low pressure turbine.

[0262] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located on a casing extending radially inwardly from a frame of the gas turbine engine.

[0263] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located upstream of a first stage of the low pressure turbine on a frame forward of the low pressure turbine.

[0264] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located in a forward section of the gas turbine engine.

[0265] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located adjacent a low pressure compressor of the gas turbine engine.

[0266] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located in the gas turbine engine so as to sense rotating bolts within the gas turbine engine.

[0267] A gas turbine engine as claimed in any preceding clause, wherein the high speed camera is located on a stator of a gearbox assembly of the gas turbine engine.

[0268] A gas turbine engine as described in any preceding clause, wherein the high speed camera comprises a plurality of high speed cameras.

[0269] A gas turbine engine as claimed in any preceding clause, wherein the high-speed camera comprises at least two high-speed cameras located at substantially the same radial position relative to a centreline axis of the gas turbine engine.

[0270] A gas turbine engine as described in any preceding clause, further comprising a damping system configured to damp the torsional vibrations in the low speed shaft based on a correlation of measured torsional vibrations and experimental torsional vibrations.

[0271] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises a mechanical damper.

[0272] A gas turbine engine as claimed in any preceding clause, wherein the damping system comprises an electrical damper.

[0273] A gas turbine engine as described in any preceding clause, wherein the damping system comprises one or more clutches coupled to the low speed shaft.

[0274] A gas turbine engine as described in any preceding clause, wherein the damping system comprises a torsional damper coupled to the low speed shaft.

[0275] A gas turbine engine according to any preceding clause, further comprising an electric machine, the damping system being configured to vary a load of the electric machine.

[0276] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0277] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0278] A method of monitoring torsional vibrations in a gas turbine engine according to any preceding clause.

[0279] A method of calculating dynamic torque in a gas turbine engine according to any preceding clause.

[0280] A method for monitoring torsion in a low-speed shaft of a gas turbine engine, the method comprising: positioning one or more sensors at target locations within the gas turbine engine; collecting data using the sensor assembly on the gas turbine engine to generate collected data; calculating the torsion of the low-speed shaft based on the collected data; generating a correlation graph between the collected data and experimental data; monitoring the torsion e-data to generate filtered data; processing the filtered data to determine the frequency and amplitude of torsional vibration of the low-speed shaft; plotting the frequency and amplitude of torsional vibration; and calculating dynamic torque based on the frequency and amplitude of the torsional vibration.

[0281] A method as described in the preceding clause, wherein the one or more sensors are one or more optical sensors.

[0282] A method as in any preceding clause, wherein the one or more sensors are one or more high speed cameras.

[0283] The method of any preceding clause, further comprising positioning a first sensor of the one or more sensors axially spaced apart from a second sensor of the one or more sensors along a longitudinal centerline axis of the gas turbine engine.

[0284] A method as defined in any preceding clause, wherein the target location is configured to sense a rotating component of the gas turbine engine.

[0285] A method as in any preceding clause, wherein the target location is a most downstream stage of the low pressure turbine.

[0286] A method as in any preceding clause, wherein the target location is a stator frame downstream of a most downstream stage of the low pressure turbine.

[0287] A method as in any preceding clause, wherein the target location is a casing extending radially inwardly from a frame of the gas turbine engine.

[0288] A method as in any preceding clause, wherein the target location is upstream of a first stage of the low pressure turbine on a frame forward of the low pressure turbine.

[0289] A method as described in any preceding clause, wherein the target location is a forward section of the gas turbine engine.

[0290] A method as described in any preceding clause, wherein the target location is a low pressure compressor of the gas turbine engine.

[0291] A method as in any preceding clause, wherein the target location is configured to sense a rotating bolt within the gas turbine engine.

[0292] A method as claimed in any preceding clause, wherein the target location is a stator of a gearbox assembly of the gas turbine engine.

[0293] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft based on the calculated dynamic torque.

[0294] A method as in any preceding clause, further comprising mechanically damping the vibrations of the low speed shaft.

[0295] A method as in any preceding clause, further comprising electrically damping said vibrations of said low speed shaft.

[0296] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft using one or more clutches coupled to the low speed shaft.

[0297] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft with a torsional damper coupled to the low speed shaft.

[0298] A method as in any preceding clause, further comprising damping the vibrations of the low speed shaft by varying a load of an electric machine coupled to the low speed shaft.

[0299] A method as claimed in any preceding clause, wherein the gas turbine engine is a non-ducted gas turbine engine.

[0300] A method as claimed in any preceding clause, wherein the gas turbine engine is a ducted gas turbine engine.

[0301] Although the above description is directed to the preferred embodiments of the present disclosure, other changes and modifications will be apparent to those skilled in the art and may be made without departing from the spirit or scope of the present disclosure. In addition, the features described in conjunction with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. A gas turbine engine having a sensor assembly configured to detect torsional vibrations, characterized in that: The gas turbine engine comprises: a fan section having a plurality of fan blades, the fan section being configured to generate an airflow through the gas turbine engine; An air flow channel, the air flow channel having: a core passage configured to direct a first portion of the airflow from an inlet to an exhaust nozzle; and a bypass passage separate from the core passage and configured to direct a second portion of the airflow around the core passage; a low-speed shaft coupled to the plurality of fan blades and configured to rotate the plurality of fan blades; a sensor assembly coupled to the gas turbine engine and configured to detect torsional vibrations in the low-speed shaft, the sensor assembly comprising a plurality of dynamic pressure sensors in the airflow passage, the plurality of dynamic pressure sensors detecting dynamic pressures of the airflow passage, the dynamic pressures being indicative of the torsional vibrations in the low-pressure shaft, wherein the sensor assembly is configured to monitor torsional vibrations in the gas turbine engine; and A damping system is configured to damp the torsional vibrations in the low speed shaft based on a correlation of the measured torsional vibrations and the experimental torsional vibrations.

2. The gas turbine engine according to claim 1, characterized in that in, The plurality of dynamic pressure sensors include at least one dynamic pressure sensor in the core channel and at least one dynamic pressure sensor in the bypass channel.

3. The gas turbine engine according to claim 1, characterized in that Further included is a low-pressure compressor and a low-pressure turbine, wherein at least one dynamic pressure sensor of the plurality of dynamic pressure sensors is located at the fan section, the low-pressure compressor, or the low-pressure turbine.

4. The gas turbine engine according to claim 1, characterized in that in, The damping system includes a mechanical damper.

5. The gas turbine engine according to claim 1, characterized in that in, The damping system includes an electrical damper.

6. The gas turbine engine according to claim 1, characterized in that in, The damping system includes one or more clutches or torsional dampers coupled to the low speed shaft.

7. The gas turbine engine according to claim 1, characterized in that Further comprising a motor, the damping system being configured to vary a load of the motor.

8. The gas turbine engine according to claim 1, characterized in that in, The plurality of dynamic pressure sensors include a first dynamic pressure sensor, a second dynamic pressure sensor, and a third dynamic pressure sensor.

9. The gas turbine engine according to claim 8, characterized in that in, The first dynamic pressure sensor is located in the bypass passage.

10. The gas turbine engine according to claim 9, characterized in that in, The first dynamic pressure sensor is located behind the fan section.

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