Gas turbine engine with sensor assembly for monitoring propeller swirl

Abstract

A gas turbine engine includes a frame, a plurality of fan blades configured to rotate about a longitudinal centerline axis of the gas turbine engine, and a mounting assembly coupling the frame to a structure. The gas turbine engine includes a sensor assembly coupled to the mounting assembly and having at least one sensor configured to detect a load on the mounting assembly and to take a corrective action when detecting propeller swirl above a predetermined limit, the feature is configured to take a corrective action when a propeller swirl is detected above a predetermined limit.

Description

Technical Field

[0001] The present disclosure generally relates to a gas turbine engine including a sensor assembly for monitoring propeller swirl. Background Art

[0002] Turbine engines generally include a fan and a core section arranged in flow communication with each other. The fan includes a plurality of fan blades that rotate about a longitudinal centerline axis of the engine. 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 diagram of a non-ducted three-flow gas turbine engine according to the present disclosure is shown, taken along the longitudinal centerline axis of the engine.

[0005] Figure 2 An exemplary sensor assembly according to the present disclosure is shown. Figure 1 A schematic perspective view of an engine.

[0006] Figure 3A Another exemplary sensor assembly according to the present disclosure is shown. Figure 1 A simplified schematic diagram of the engine.

[0007] Figure 3B It shows that according to the present disclosure Figure 3A Simplified front and rear view of the front bearing assembly.

[0008] Figure 3C It shows that according to the present disclosure Figure 3A Simplified front and rear view of the rear bearing assembly.

[0009] Figure 4A A fan blade (such as Figure 1 Schematic diagram of the forces on the fan blades of a turbine engine shown.

[0010] Figure 4B A fan blade (such as Figure 1 Schematic diagram of the forces on the fan blades of a turbine engine shown.

[0011] Figure 5 A method for detecting fan blades (such as Figure 1 Method of swirl of fan blades of a turbine engine as shown.

[0012] Fig. 6AA schematic perspective view of a non-ducted engine with controllable fan blades according to the present disclosure is shown.

[0013] Figure 6B The present disclosure shows Fig. 6A Schematic diagram of the phase difference between the fan blades of a non-ducted engine.

[0014] Figure 7 The invention shows a load reduction device according to the present invention. Figure 1 A simplified schematic diagram of the engine.

[0015] Figure 8 An exemplary display of a system for detecting swirl of a fan blade according to the present disclosure is shown. DETAILED DESCRIPTION

[0016] Features, advantages and embodiments of the present disclosure are set forth or apparent by considering the following detailed description, drawings and claims.In addition, the following detailed description is exemplary and is intended to provide further explanation without limiting the scope of the present disclosure as claimed.

[0017] 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.

[0018] 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.

[0019] 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.

[0020] 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.

[0021] 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.

[0022] 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.

[0023] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0024] 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.

[0025] 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.

[0026] 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.

[0027] The present disclosure describes a detection engine (such as Figure 1 A system and method for swirl in a fan blade of a gas turbine engine (described herein). As used herein, a fan blade refers to an airfoil extending radially from a center disk in a fan section or propeller section of a turbine engine. A fan blade is any rotating airfoil used to generate thrust in ducted engines and non-ducted engines. Therefore, the term fan blade includes not only fan blades, but also propellers and propeller blades. Swirl is also referred to herein as propeller swirl, propeller swirl, swirl mode, propeller swirl mode, and propeller swirl mode. During operation of the engine, as will be described in more detail below, the fan blades may experience propeller swirl, which may cause the fan blades to rotate off-center from the center axis of the engine, change the angle of attack of the fan blades, or both. This alternation of the operation of the fan blades affects the performance of the engine. The present disclosure provides a sensor assembly for measuring the load on a mounting link that couples an engine to an aircraft. The present disclosure provides a sensor assembly for measuring the load on a bearing assembly associated with a low-pressure shaft.

[0028] That is, the present disclosure provides strain gauges for calculating individual mounting loads. The mounting loads, along with the mounting geometry, can be used to calculate engine operating loads. The engine operating loads can be further divided into normal or nominal engine operation, maneuvering loads (e.g., aerodynamic loads), and dynamic loads that capture IP loads (e.g., loads at base RPM of the fan blades). The relationship between the mounting loads and maneuvering loads can be obtained through a set of linear equations and / or a transfer function can be developed using development engine testing.

[0029] 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.

[0030] like Figure 1As 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.

[0031] 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 1 As 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.

[0032] 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.

[0033] 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 1 Schematically 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.

[0034] 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.

[0035] 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 1In 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.

[0036] 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.

[0037] 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.

[0038] The fan assembly 50 also includes an intermediate fan 86. The intermediate fan 86 includes a plurality of intermediate fan blades 88 ( Figure 1Only 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.

[0039] 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.

[0040] 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.

[0041] 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.

[0042] 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 33 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 .

[0043] 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.

[0044] 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.

[0045] Figure 1 The gas turbine engine 10 depicted in FIG. 1 is provided as an example only. In other embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, in other embodiments, the primary fan 52 may be configured in any other suitable manner (e.g., as a fixed pitch fan), and any other suitable fan frame (e.g., Figure 2 The fan frame 98 of the present invention is supported by the structure of the fan frame 98. In other embodiments, the primary fan 52 can be ducted through the fan case or the 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 still other embodiments, aspects of the present disclosure can be incorporated into any other suitable turbine engine (e.g., a turbofan engine, a propfan 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)).

[0046] 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.

[0047] 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 1 The 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.

[0048] Brief reference Figure 4A and Figure 4B During operation, fan blades 54 rotate in a clockwise direction 408 about longitudinal centerline axis 12 while airflow A flows from the front end of the engine to the rear end of the engine (omitted for clarity). Figure 4A and Figure 4B A single representative fan blade 54 is shown to facilitate understanding of the propeller swirl pattern. The fan blade 54 rotates three hundred and sixty degrees from the twelve o'clock position 400 to the three o'clock position 402, to the six o'clock position 404, to the nine o'clock position 406, and back to the twelve o'clock position 400. The fan blade 54 is a rotating mass about the longitudinal centerline axis 12, with the constrained end at the base of the fan blade 54. Therefore, when the fan blade 54 rotates about the longitudinal centerline axis 12, the load exists at the free end of the fan blade 54 (e.g., closer to the tip of the fan blade 54 than the constrained base of the fan blade 54) in the axial direction 410 and the vertical direction 412.

[0049] Thus, during operation at the 1P frequency (eg, the base RPM of the fan blade 54), the rotating fan blade 54 will experience two phenomena depending on the position of the fan blade 54, such as Figure 4A and Figure 4B As shown. Figure 4A , fan blade 54 experiences shear and a downward moment in vertical direction 450, which causes fan blade 54 to yaw to the right in yaw direction 460 (e.g., twist or oscillate about vertical direction 450). The vertically directed shear and right-directed yaw cause fan blade 54 to pitch upward. Figure 4B , fan blades 54 also experience shear and lateral movement in a lateral direction 470, which causes fan blades 54 to move in a pitch direction 480 (e.g., in a direction corresponding to the pitch direction 480). Figure 4A The fan blade 54 is pitched downwardly (in a direction opposite to the vertical direction 450 of the fan blade). The right-directed shear and the downward-directed pitch cause the fan blade 54 to yaw (e.g., twist or oscillate) to the right. The above-described phenomenon experienced by the rotating fan blade 54 represents the propeller swirl of the fan blade 54. The propeller swirl at the 1P frequency excites the backward swirl motion 490 in the fan blade 54.

[0050] In other words, when the fan blade 54 rotates about the longitudinal centerline axis 12, the fan blade 54 has different angles of attack at the various "clock" positions described above. That is, the fan blade 54 experiences different amounts of lift. When installed in an aircraft, this may occur when the aircraft takes off, and there is a takeoff rotation (e.g., 1P frequency) in the aircraft. The change in the angle of attack of the fan blade 54 causes a moment in the yaw direction (e.g., direction 460) and the pitch direction (e.g., pitch direction 480). The yaw moment is fed into the pitch moment, and the pitch moment response is fed into the yaw moment. This causes a propeller swirl, which causes large yaw and pitch moments on the fan rotor system. The fan rotor system is mounted to the aircraft through a plurality of mounting rods, which are loaded due to the rotation of the fan blade 54. Therefore, the sensor assembly of the present disclosure detects the propeller swirl by measuring the force acting on the mounting rods that connect the engine 10 to the aircraft (omitted for clarity).

[0051] Return reference, Figure 2 Shows Figure 1 10. Although for clarity, Figure 1Although omitted in the drawings, the engine includes a turbine aft frame 97 and a fan frame 98. The fan frame 98 houses the fan rotor system. A mounting assembly 100 couples the turbine aft frame 97 and the fan frame 98 to the aircraft (omitted for clarity). The mounting assembly 100 includes a front mount 101 having a first front mount link 102 and a second front mount link 104. The mounting assembly 100 includes a rear mount 106 having a first rear mount link 108 and a second rear mount link 110. The mounting assembly 100 includes a thrust mount link 112. The front mount 101 couples the fan frame 98 to the aircraft, and the rear mount 106 couples the turbine aft frame 97 to the aircraft. The thrust mount link 112 couples the turbine aft frame 97 to the fan frame 98 via a thrust mount 114.

[0052] Continue to refer Figure 2 , the engine 10 includes a sensor assembly 200. The sensor assembly 200 includes one or more sensors coupled to the mount assembly 100. The sensor assembly 200 may include a first front sensor 202 on the first front mount link 102. The sensor assembly 200 may include a second front sensor 204 on the second front mount link 104. The sensor assembly 200 may include a first rear sensor 206 on the first rear mount link 108. The sensor assembly 200 may include a second rear sensor 208 on the second rear mount link 110. The sensor assembly 200 may include a thrust mount sensor 210 coupled to the thrust mount link 112.

[0053] The sensors of sensor assembly 200 may be strain gauges or other sensors configured to detect loads (e.g., reaction loads) on the mounting member links of mounting assembly 100. Each of the strain gauges is oriented along the length or longitudinal direction of each mounting member link. The mounting member links of mounting assembly 100 are loaded along the length or longitudinal direction, so that the mounting member links are in tension or compression. Therefore, the strain gauges are oriented in the same direction as the direction in which the mounting member links are loaded. Although five sensors are shown and described, more or fewer sensors may be provided.

[0054] As previously discussed, mounting assembly 100 couples engine 10 to the aircraft. Thus, since the sensors of sensor assembly 200 are associated with the mounting rods of mounting assembly 100, any loads generated on engine 10 will be reacted in mounting assembly 100 and will be recorded by sensor assembly 200. Thus, each sensor indicates the overall reaction of engine 10 (e.g., by measuring the reaction loads), including any pitch and yaw moments discussed with respect to FIG. 4. As will be described with reference to FIG. Figure 56 , loads unrelated to pitch and yaw moments indicative of propeller swirl may be filtered out of the signals transmitted from sensor assembly 200 . Thus, the strain gauges of sensor assembly 200 detect propeller swirl of fan blade 54 .

[0055] FIG. 3A to FIG. 3C An alternative sensor assembly 300 is shown that may be coupled to engine 10. Sensor assembly 300 may be used in place of sensor assembly 200 or in conjunction with sensor assembly 200. FIG. 3A to FIG. 3C Similar reference numerals in Figure 1 The sensor assembly 300 is coupled to the bearing assembly, which is coupled to the low-pressure shaft 38. The sensor assembly 300 may include a first front sensor 302 coupled to the front bearing assembly 116. The sensor assembly 300 may include a second front sensor 304 coupled to the front bearing assembly 116. The first front sensor 302 and the second front sensor 304 may be mounted to the front bearing assembly 116 in an orientation ninety degrees apart, such as Figure 3B 300 is shown in a front-to-rear orientation, wherein the first front sensor 302 is coupled at the twelve o'clock position and the second front sensor 304 is coupled at the three o'clock position. This allows the sensor assembly 300 to obtain the load at the front end of the engine 10 in the vertical and lateral directions.

[0056] The sensor assembly 300 may include a first rear sensor 306 coupled to the rear bearing assembly 118. The sensor assembly 300 may include a second rear sensor 308 coupled to the rear bearing assembly 118. The first rear sensor 306 and the second rear sensor 308 may be mounted to the rear bearing assembly 118 in an orientation ninety degrees apart, such as Figure 3C 300 is shown in a front-to-rear orientation, wherein the first rear sensor 306 is coupled at the twelve o'clock position and the second rear sensor 308 is coupled at the three o'clock position. This allows the sensor assembly 300 to obtain the load at the rear end of the engine 10 in both the vertical and lateral directions.

[0057] The first front sensor 302, the second front sensor 304, the first rear sensor 306, and the second rear sensor 308 may be located on the stator of the bearing assembly, respectively. Although four sensors are shown and described, more or fewer sensors may be provided. For example, only two front sensors (e.g., the first front sensor 302 and the second front sensor 304) may be provided.

[0058] Although the sensor assembly 200 ( Figure 2 ) and sensor assembly 300( Figure 3A) are shown and described as including the above-described sensors, but any combination or sub-combination of sensors may be present, and / or there may be fewer sensors than those shown and described. The sensors of sensor assembly 200 and sensor assembly 300 are described herein as strain gauges. Other sensor types (such as load sensors and / or accelerometers) may be used in place of or in combination with any one or more sensors of sensor assembly 200 and sensor assembly 300.

[0059] refer to Figure 5 , the method 500 can detect a sensor assembly 200 ( Figure 2 ) or sensor assembly 300( Figure 3A ) of a gas turbine engine 10( Figure 1 ) in the fan blades 54. At step 502, the engine 10, and therefore the fan blades 54, are placed into operation (e.g., rotation begins), and aerodynamic forces act on each fan blade 54 as previously described. At step 504, sensors of the sensor assembly 200 or the sensor assembly 300 detect and transmit data. At step 506, the measured data is used to calculate the dynamic propeller swirl of the fan blades 54. At step 508, the calculation is compared with a predetermined limit. The predetermined limit may represent a value when the propeller swirl is detected. In some examples, the predetermined limit may represent a maximum acceptable value for the propeller swirl. If the calculated propeller swirl is greater than the predetermined limit, action may be taken at step 510, such as a remedial action or a corrective action. After taking the action at step 510 and / or if the calculated propeller swirl is less than the predetermined limit, the method 500 is cyclically repeated from step 504 to step 510 until stopping of the engine 10 occurs at step 512. That is, the method 500 is performed when the engine 10 ( Figure 1 ) is a closed loop that operates continuously and in real time during operation of the . Although described as predetermined limits, the calculated propeller swirl may be compared to a predetermined range of acceptable values.

[0060] The action taken at step 510 can be an active action, a passive action, or a combination thereof. Exemplary active actions can include pitch control of a fan blade (e.g., fan blade 54 described with respect to Fig. 3). For example, each fan blade can be controlled independently of each other fan blade by a pitch control system. In another example, each fan blade can be controlled at a specific clock position around the axis of rotation. The pitch of the fan blade can be controlled during each rotation of each fan blade. The fan blade can be controlled based on the angular position to control the swirl torque, thereby reducing the propeller swirl.

[0061] Fig. 6A The method for performing step 510 ( Figure 5) is an exemplary engine 600 with multiple independently controllable fan blades 602 that are actively actuated. Each fan blade 602 has a blade axis 604, and the fan blade 602 has a pitch angle around the blade axis 604. Each fan blade 602 is controlled relative to the blade axis 604 to orient the fan blade 602 at a specific pitch angle relative to the blade axis 602. Each fan blade 602 is controlled by rotating the fan blade 602 around the blade axis 604 in a direction 606. The pitch angle of each fan blade 602 can be controlled independently of the remaining fan blades 602. For example, a first fan blade 602a can be controlled independently of the other fan blades around the first blade axis 604a, and similarly, the same is true for the remaining fan blades 602.

[0062] The pitch angle of each fan blade 602 affects the forces acting on fan blade 602 (e.g., as Figure 4A and Figure 4B The change in pitch angle affects the yaw angle of the fan blade 602, thereby affecting the propeller swirl, as described above. Figure 4A and Figure 4B and are described by the following relations (1) to (6). In an example, each fan blade 602 can be controlled to be out of phase with each other fan blade 602. For example, referring to Figure 6B , the first fan blade 602a and the second fan blade 602b are shown as a function of the clock position of the blades. That is, for example, referring to Figure 4A , Figure 4B and Fig. 6A , each fan blade rotates three hundred and sixty degrees about the longitudinal engine centerline axis. Each position of the fan blade through these three hundred and sixty degrees can be associated with a clock position. For example, zero degrees can be associated with the twelve o'clock position (e.g., Figure 4A 400 in ), one hundred and eighty degrees may be associated with the six o'clock position (e.g., Figure 4A 404). As the fan blades rotate about the engine centerline axis, changes in the pitch angle affect the relative position of the fan blades relative to the engine centerline axis.

[0063] As described above, when at step 510 ( Figure 5 ) When active action is adopted, each fan blade 602 is controlled independently of the other fan blades 602. Figure 6B In the example of FIG. 1 , each fan blade 602 is controlled to be out of phase with the remaining fan blades. That is, if the first fan blade 602a is at the first pitch angle, when the first fan blade 602a rotates about the engine centerline axis, the pitch angle will be according to Figure 6BThe second fan blade 602b may be controlled to be at a second pitch angle such that when the second fan blade 602b rotates about the engine centerline axis, the sine wave of the second fan blade 602b is misaligned or out of phase with the first fan blade 602a. The same is true for the remaining fan blades. In this way, each pitch angle of the fan blades is misaligned to reduce the forces acting on the blades, thereby reducing or reducing the propeller swirl. In one example, the fan blades may be out of phase in a uniform distribution of approximately three hundred and sixty degrees. For example, in a fan having five fan blades 602 Fig. 6A In the example of FIG. 5 , each fan blade 602 may be seventy-two degrees out of phase with respect to adjacent fan blades 602 .

[0064] An example passive action may be a load reduction device. A passive action may be a fault protection. Figure 7 An exemplary load reduction device 700 is shown in FIG. Figure 7 Only engine 10 ( Figure 1 ). That is, in Figure 7 Only the fan 52 and the low speed shaft 38 are shown, however, Figure 1 All the characteristics of Figure 2 and either or both of the sensor systems of FIG. 3 may be present in Figure 7 That is, the load reduction device 700 can be employed in any of the embodiments described herein.

[0065] refer to Figure 7 , the load reduction device 700 can be included in a bearing assembly 716. The bearing assembly 716 includes a plurality of bearings 702 and a bearing housing 704. The bearing housing 704 has a cylindrical portion 706 and a tapered portion 708. The tapered portion 708 can include a region 710 having a reduced thickness compared to the remainder of the tapered portion 708, such that when the load on the bearing housing 704 exceeds a predetermined value, the region 710 bends, breaks, or otherwise fails.

[0066] During operation, when the load on the bearing housing 704 exceeds a predetermined value, the load reduction device 700 is configured to bend or fail at the region 710. When the predetermined value associated with the propeller swirl is above a predetermined critical limit (e.g., Figure 8 When the load reduction device 700 is bent or fails to protect the components connected to the fan blades, the engine is then shut down once the load reduction device 700 fails.

[0067] Return to reference Figure 5, the calculation of step 506 converts the strain / load data collected from the sensor assembly 200 into an indication of propeller swirl. The calculation of step 506 divides the total sensed installation load into (1) normal operating loads experienced by the mounting assembly 100 and (2) installation loads associated with propeller swirl. The load data collected by the sensor assembly 200 is represented by the following relationship 1.

[0068] {f} total = {f} nominal +{f} inlet aero +{f} 1P load (1) where {f} total represents the total load experienced by the mounting assembly 100, {f} nominal represents the nominal engine operating load, {f} inlet aero represents the aerodynamic load and is based on the engine test conditions, and {f} IP load represents the loads experienced by the mounting assembly 100 related to the propeller swirl and the 1P load. Each of the above loads is related to the geometry of the mounting assembly 100. Each of the above loads can be represented as a factor matrix as shown in (2) to (5) below.

[0069] As shown in equation (2), each load (including the loads in equations (3), (4) and (5)) can be represented by three lateral forces (e.g., axial force, vertical force and lateral force) and three moments (e.g., roll, pitch and yaw).

[0070]

[0071] As shown in relation (3), the nominal operating load includes the axial force due to the thrust or acceleration of the engine, the vertical force due to gravity, and the pitch moment and yaw moment due to the gyroscopic load generated when the rotation axis of the rotor of the engine changes. Compared with the 1P load, the load due to the nominal operating load is generally a more stable state load.

[0072]

[0073] As shown in equation (4), the aerodynamic loads include vertical force due to vertical shear force, lateral force due to lateral shear force, pitch moment and yaw moment. Compared with 1P loads, aerodynamic loads are usually more stable state loads.

[0074]

[0075] As shown in relation (5), the 1P load includes the vertical force due to vertical shear, the lateral force due to lateral shear, the pitch moment and the yaw moment. The IP load is a dynamic load that changes during operation of the engine.

[0076]

[0077] According to the above matrix, the pitch moment and load moment described in FIG. 4 (e.g., dynamic load due to 1P load, e.g., {f} IP load ) can be obtained by taking the total sensing data {f} total Subtract or remove the nominal load {f} nominal and aerodynamic loads {f} inlet aero Once determined, the dynamic load {f} IP load The data obtained from the sensor assembly 200 is equal to {f} as described above. total . {f} inlet aero and {f} nominal Both can be determined from engine test data. Therefore, the {f} indicating propeller swirl IP load It can be determined from the sensor assembly data by removing or subtracting the engine test data associated with the above loads, as shown in relation (6). This is achieved by passing the strain gauge data from the sensor assembly 200 through a band pass filter to remove the nominal load and aerodynamic load from the total strain gauge data.

[0078] {f} 1P load ={f} total - {f} nominal - {f} inlet aero (6)

[0079] Although the above description of step 506 is described with respect to sensor assembly 200 , sensor assembly 300 may employ similar or identical processing.

[0080] Figure 8 It is shown that the sensor assembly 200 ( Figure 2 ) or sensor assembly 300( Figure 3A ) and Method 500( Figure 5) is an exemplary display 800 of data obtained by a load sensor assembly 200 or sensor assembly 300. Display 800a shows a graph of load detected by sensor assembly 200 or sensor assembly 300 relative to the time or duration of detection. As shown, data 801 from sensor assembly 200 or sensor assembly 300 is plotted as a function of the time or duration of detection.

[0081] The data 801 may be below the first predetermined limit 802, which may indicate an acceptable or "green" load on the mounting assembly 100 (or on the bearing assembly of the sensor assembly 300). In the event that the data 801 is below the first predetermined limit 802, no action is taken (e.g., step 510), and monitoring (e.g., steps 504, 506, 508) is continued.

[0082] Data 801 may sometimes extend beyond first predetermined limit 802, such that data 801 may be positioned above first predetermined limit 801, but below second predetermined limit 804. This may represent a warning or "yellow" load on mounting assembly 100 (or on a bearing assembly of sensor assembly 300). Second predetermined limit 804 is an unacceptable load value on a fan blade that may cause a failure within the engine. When data 801 is within this yellow section (e.g., between first predetermined limit 802 and second predetermined limit 804), this is considered an alarm or warning zone. Monitoring continues within the yellow section. Active action (e.g., control of the blades) may occur in the yellow section.

[0083] In some cases, data 801 may be positioned above a second predetermined limit 804, also referred to as a critical limit 804, which may indicate a critical or "red" load on the mounting assembly (or on a bearing assembly of sensor assembly 300). When critical limit 804 is reached, a passive action (e.g., a load reduction device) may be actuated to reduce the moment on the fan blades and reduce propeller swirl.

[0084] Each of the first predetermined limit 802 and the second predetermined limit 804 can be set or predetermined based on a specific engine, specific expected operating conditions, a specific aircraft, the architecture of the engine, the engine size, the fan size, etc., or a combination thereof. Display 800b shows data 801 processed using a bandpass filter 806. Once processed, data 801 becomes data 803 and is plotted as a relationship between 1P load (e.g., the load at the base RPM of the fan blade 54) and frequency. Display 800b indicates whether there is a propeller swirl in the fan blade 54.

[0085] The green or acceptable range (e.g., below the first predetermined limit 802) can indicate a load that is acceptable for operation and no remedial action is required. The yellow or warning range (e.g., between the first predetermined limit 802 and the second predetermined limit 804) can indicate a situation where the operator monitors or otherwise observes to ensure that the second predetermined limit 804 is not exceeded. Remedial action may or may not be taken in the yellow range. In the red or critical range (e.g., above the second predetermined limit 804), remedial action may then be taken to reduce the load to below the second predetermined limit 804 and / or pause or stop the operation of the engine 10.

[0086] Therefore, the present disclosure describes systems and methods for detecting swirl in the fan blades of an engine. The sensor assembly and related methods of the present disclosure provide an engine monitoring system. The sensor assembly and related methods of the present disclosure allow for detection of propeller swirl in an open fan engine and detection of related 1P loads within defined limits relative to the flight mission. The present disclosure allows for conversion of strain measurements at a mounting rod or bearing assembly into detection of propeller swirl. The present disclosure is based on a combination of mounting rod load measurements using strain gauges and their correlation to 1P dynamic loads using free body diagrams and / or development of engine traction load test correlations. The present disclosure also provides for direct instrumentation of low pressure shaft bearings with strain gauges and / or accelerometers to measure dynamic content from 1P loads.

[0087] Further aspects of the disclosure are provided by the subject matter of the following clauses.

[0088] A gas turbine engine having a sensor assembly for monitoring propeller swirl, the gas turbine engine comprising: a frame that houses the gas turbine engine; a plurality of fan blades that are configured to rotate about a longitudinal centerline axis of the gas turbine engine; a mounting assembly that couples the frame to a structure; a sensor assembly that is coupled to the mounting assembly and has at least one sensor that is configured to detect a load on the mounting assembly, the sensor assembly being configured to monitor propeller swirl in the plurality of fan blades by detecting the load on the mounting assembly; and a feature that is configured to take corrective action when propeller swirl is detected to be above a predetermined limit.

[0089] The gas turbine engine according to the preceding clause, wherein said at least one sensor is a strain gauge.

[0090] A gas turbine engine as described in any preceding clause, wherein said at least one sensor is a load sensor.

[0091] According to the gas turbine engine described in any preceding item, the mounting assembly includes a front mount, a thrust mount connecting rod and an aft mount, and the at least one sensor includes a first front sensor and a second front sensor on the front mount, a first aft sensor and a second aft sensor on the aft mount, and a thrust sensor on the thrust mount connecting rod.

[0092] A gas turbine engine as described in any preceding clause, wherein the mounting assembly includes a front mount, the at least one sensor being coupled to the front mount.

[0093] A gas turbine engine as in any preceding clause, said mounting assembly comprising a front mount, said at least one sensor comprising a first front sensor and a second front sensor each coupled to said front mount.

[0094] A gas turbine engine as described in any preceding clause, wherein the mounting assembly comprises a thrust mount connecting rod, the at least one sensor being coupled to the thrust mount connecting rod.

[0095] A gas turbine engine as described in any preceding clause, wherein the mounting assembly comprises a thrust mount connecting rod, the at least one sensor comprising a thrust sensor coupled to the thrust mount connecting rod.

[0096] A gas turbine engine as described in any preceding clause, said mounting assembly comprising an aft mounting, said at least one sensor being coupled to said aft mounting.

[0097] A gas turbine engine as in any preceding clause, said mounting assembly comprising an aft mount, said at least one sensor comprising a first aft sensor and a second aft sensor each coupled to said aft mount.

[0098] A gas turbine engine as claimed in any preceding clause, wherein said feature is a load reduction device configured to fail when said propeller swirl is detected to be above said predetermined limit.

[0099] A gas turbine engine as described in any preceding clause, wherein said feature is means for controlling the pitch of each of said plurality of fan blades when said propeller swirl is detected to be above said predetermined limit.

[0100] The gas turbine engine of any preceding clause, said frame comprising a fan frame, and said mounting assembly comprising a front mount coupling said fan frame to said structure, and said at least one sensor being located on said front mount.

[0101] According to the gas turbine engine of any preceding clause, the front mount includes a first front mount link and a second front mount link, and the at least one sensor includes a first front sensor located on the first front mount link and a second front sensor located on the second front mount link.

[0102] A gas turbine engine as in any preceding clause, said frame comprising a turbine aft frame, and said mounting assembly comprising an aft mount coupling said turbine aft frame to said structure, and said at least one sensor comprising a sensor located on said aft mount.

[0103] According to the gas turbine engine of any preceding clause, the aft mount includes a first aft mount connecting rod and a second aft mount connecting rod, and the at least one sensor includes a first aft sensor located on the first aft mount connecting rod and a second aft sensor located on the second aft mount connecting rod.

[0104] The gas turbine engine of any preceding clause, further comprising a thrust mount connecting rod coupled to the aft mount, the at least one sensor comprising a thrust mount sensor located on the thrust mount connecting rod.

[0105] A gas turbine engine according to any preceding clause, wherein the frame comprises a fan frame and a turbine aft frame, and wherein the at least one sensor comprises a forward sensor located on a forward mount of the mounting assembly and an aft sensor located on an aft mount of the mounting assembly.

[0106] According to the gas turbine engine described in any preceding item, the front sensor includes a first front sensor and a second front sensor, the first front sensor is on a first front mount connecting rod connecting the front mount to the fan frame, and the second front sensor is on a second front mount connecting rod connecting the front mount to the fan frame.

[0107] According to the gas turbine engine described in any preceding item, the rear sensor includes a first rear sensor and a second rear sensor, the first rear sensor is on a first rear mounting rod connecting the rear mounting member to the turbine rear frame, and the second rear sensor is on a second rear mounting rod connecting the rear mounting member to the turbine rear frame.

[0108] A gas turbine engine according to any preceding clause, further comprising a thrust sensor coupled to a thrust mount of the turbine aft frame.

[0109] An aircraft comprising a gas turbine engine according to any preceding clause.

[0110] An aircraft, comprising: a non-ducted engine, the non-ducted engine including a longitudinal centerline axis and a plurality of fan blades configured to rotate about the longitudinal centerline axis; a mounting assembly, the mounting assembly coupling the non-ducted engine to the aircraft; a sensor assembly coupled to the mounting assembly and configured to detect a load on the mounting assembly, the sensor assembly configured to monitor propeller swirl in the plurality of fan blades by detecting the load on the mounting assembly; and a feature, the feature configured to take corrective action when propeller swirl is detected to exceed a predetermined limit.

[0111] An aircraft as described in any preceding clause, wherein at least one sensor of the sensor assembly is a strain gauge or a load sensor.

[0112] An aircraft according to any preceding clause, wherein the mounting assembly includes a front mount, a thrust mount link and a rear mount, and the sensor assembly includes first and second front sensors on the front mount, first and second rear sensors on the rear mount, and a thrust sensor on the thrust mount link.

[0113] An aircraft as described in any preceding clause, wherein the feature is a load reduction device, a device configured to control the pitch of the plurality of fan blades, or both the load reduction device and the device configured to control the pitch of the plurality of fan blades.

[0114] An aircraft as described in any preceding clause, further comprising a frame between the non-ducted engine and the aircraft.

[0115] An aircraft as in any preceding clause, said frame comprising a fan frame and a turbine aft frame, and said sensor assembly comprising a forward sensor located on a forward mount of said mounting assembly and a rearward sensor located on a rearward mount of said mounting assembly.

[0116] An aircraft according to any preceding clause, wherein the front sensor includes a first front sensor and a second front sensor, wherein the first front sensor is on a first front mount link connecting the front mount to the fan frame, and the second front sensor is on a second front mount link connecting the front mount to the fan frame.

[0117] An aircraft according to any preceding clause, wherein the rear sensor includes a first rear sensor and a second rear sensor, wherein the first rear sensor is on a first rear mount link coupling the rear mount to the turbine rear frame, and the second rear sensor is on a second rear mount link coupling the rear mount to the turbine rear frame.

[0118] An aircraft as described in any preceding clause, further comprising a thrust sensor coupled to a thrust mount of the turbine aft frame.

[0119] An aircraft as in any preceding clause, said frame comprising a fan frame, and said mounting assembly comprising a front mount configured to couple said fan frame to said structure, and said at least one sensor being located on said front mount.

[0120] An aircraft as in any preceding clause, the front mount comprising a first front mount link and a second front mount link, and the at least one sensor comprising a first front sensor located on the first front mount link and a second front sensor located on the second front mount link.

[0121] An aircraft as in any preceding clause, said frame comprising a turbine aft frame, and said mounting assembly comprising an aft mount configured to couple said turbine aft frame to said structure, and said at least one sensor comprising a sensor located on said aft mount.

[0122] An aircraft as described in any preceding clause, the rear mount comprising a first rear mount link and a second rear mount link, and the at least one sensor comprising a first rear sensor located on the first rear mount link and a second rear sensor located on the second rear mount link.

[0123] An aircraft as described in any preceding clause, further comprising a thrust mount link coupled to the aft mount, the at least one sensor comprising a thrust mount sensor located on the thrust mount link.

[0124] An aircraft as claimed in any preceding clause, further comprising a load reduction device configured to fail upon detecting that the propeller swirl is above the predetermined limit.

[0125] An aircraft as claimed in any preceding clause, further comprising means for controlling the pitch of each of said plurality of fan blades upon detecting that said propeller swirl is above said predetermined limit.

[0126] A sensor assembly for detecting propeller swirl in a fan blade, the sensor assembly comprising a first front sensor, a second front sensor, a thrust sensor, a first rear sensor and a second rear sensor, each of the sensors being a strain gauge or a load sensor.

[0127] A gas turbine engine having a sensor assembly according to any preceding clause.

[0128] A method of detecting propeller swirl using a sensor assembly according to the preceding clause.

[0129] A method of detecting propeller swirl of a fan blade of a gas turbine engine according to any preceding clause.

[0130] A method of detecting propeller swirl of a fan blade of an aircraft according to any preceding clause.

[0131] A method of detecting propeller swirl of a fan blade of an engine, the method comprising operating the engine, rotating the fan blade in a circumferential direction during operation of the engine, detecting loads on a mounting assembly configured to retain the engine on a structure using a sensor, filtering the detected loads to remove nominal and aerodynamic loads to generate a 1P load, comparing the 1P load to a predetermined limit, and detecting propeller swirl of the fan blade based on the comparison.

[0132] A method as recited in the preceding clause, further comprising taking remedial action when propeller swirl is detected.

[0133] A method as in any preceding clause, filtering the detected load further comprising filtering using a bandpass filter.

[0134] A method as in any preceding clause, wherein the sensor is a strain gauge.

[0135] A method as in any preceding clause, wherein the engine is a non-ducted engine.

[0136] A method as in any preceding clause, further comprising taking action when the IP load is above the predetermined limit.

[0137] A method as in any preceding clause, wherein the action is a passive action comprising failure of a load reduction device when the propeller swirl is above the predetermined limit.

[0138] A method as in any preceding clause, wherein the action is an active action comprising controlling the pitch of each of the plurality of fan blades when the propeller swirl is above the predetermined limit.

[0139] A method as in any preceding clause, wherein the detected load is a total load experienced by the mounted assembly, the total load being equal to the sum of a nominal engine operating load, an aerodynamic load and the IP load.

[0140] A method as in any preceding clause, wherein filtering the sensed load comprises subtracting each of the nominal engine operating load and the aerodynamic load from the total load to produce the IP load.

[0141] A method as in any preceding clause, wherein each of the total load, the nominal engine operating load, the aerodynamic load and the IP load comprises an axial force, a vertical force, a lateral force, a rolling moment, a pitch moment and a yaw moment.

[0142] A method as in any preceding clause, wherein the nominal engine loads include axial forces due to thrust, vertical forces due to gravity, pitch moments due to rotational motion, and yaw moments.

[0143] A method as defined in the preceding clause, wherein the lateral force and the rolling moment at said nominal engine load are both zero.

[0144] A method as in any preceding clause, wherein the aerodynamic loads include vertical and lateral forces each due to shear, a pitch moment and a yaw moment due to aerodynamics.

[0145] A method as defined in the preceding clause, wherein the aerodynamic load comprises zero axial force and zero rolling moment.

[0146] A method as in any preceding clause, wherein the IP loads include vertical and lateral forces each due to shear, a pitch moment and a yaw moment due to IP dynamic loads.

[0147] A method as defined in the preceding clause, wherein the aerodynamic load comprises zero axial force and zero rolling moment.

[0148] A method as in any preceding clause, wherein the aerodynamic load and the nominal load are determined from engine test data.

[0149] An engine comprises: a plurality of fan blades, the plurality of fan blades being configured to rotate about a longitudinal centerline axis of the engine; a shaft, the shaft being configured to rotate the plurality of fan blades; a front bearing assembly, the front bearing assembly supporting a front end of the shaft; a rear bearing assembly, the rear bearing assembly supporting a rear end of the shaft; a sensor assembly, the sensor assembly being coupled to the front bearing assembly and being configured to detect a load on the front bearing assembly, the sensor assembly being configured to monitor propeller swirl in the plurality of fan blades by detecting the load on the front bearing assembly; and a feature, the feature being configured to take corrective action when propeller swirl is detected to be above a predetermined limit.

[0150] An engine as recited in the preceding clause, wherein the sensor assembly is coupled to the rear bearing assembly and is configured to detect a load on the rear bearing assembly.

[0151] An engine as claimed in any preceding clause, the sensor assembly comprising first and second front sensors each coupled to the front bearing assembly.

[0152] An engine as described in any preceding clause, the first front sensor being coupled to the front bearing assembly at the twelve o'clock position and the second front sensor being coupled to the front bearing assembly at the three o'clock position.

[0153] An engine as claimed in any preceding clause, the sensor assembly comprising first and second rear sensors each coupled to the rear bearing assembly.

[0154] An engine as described in any preceding clause, the first rear sensor being coupled to the rear bearing assembly at the twelve o'clock position and the second rear sensor being coupled to the rear bearing assembly at the three o'clock position.

[0155] An engine as described in any preceding clause, wherein the sensor assembly includes one or more strain gauges or accelerometers.

[0156] An engine as described in any preceding clause, wherein the shaft is a low pressure turbine shaft.

[0157] An engine as described in any preceding clause, further comprising a load reduction device configured to fail when said propeller swirl is detected to be above said predetermined limit.

[0158] An engine as described in any preceding clause, further comprising means for controlling the pitch of each of said plurality of fan blades when said propeller swirl is detected to be above said predetermined limit.

[0159] An engine as claimed in any preceding clause, wherein said feature is a load reduction device configured to fail upon detection of said propeller swirl above said predetermined limit.

[0160] An engine as described in any preceding clause, wherein said feature is means for controlling the pitch of each of said plurality of fan blades when said propeller swirl is detected to be above said predetermined limit.

[0161] An aircraft including a non-ducted engine, the non-ducted engine comprising: a longitudinal centerline axis and a plurality of fan blades configured to rotate about the longitudinal centerline axis; a shaft configured to rotate the plurality of fan blades; a front bearing assembly, the front bearing assembly supporting a front end of the shaft; a rear bearing assembly, the rear bearing assembly supporting a rear end of the shaft; a sensor assembly, the sensor assembly coupled to the front bearing assembly and configured to detect a load on the front bearing assembly, the sensing assembly configured to monitor propeller swirl in the plurality of fan blades by detecting the load on the front bearing assembly; and a feature configured to take corrective action when propeller swirl is detected to exceed a predetermined limit.

[0162] An aircraft as recited in the preceding clause, wherein the sensor assembly is coupled to the rear bearing assembly and is configured to detect a load on the rear bearing assembly.

[0163] An aircraft as in any preceding clause, the sensor assembly comprising a first front sensor and a second front sensor each coupled to the front bearing assembly.

[0164] An aircraft as in any preceding clause, the first front sensor coupled to the front bearing assembly at a twelve o'clock position and the second front sensor coupled to the front bearing assembly at a three o'clock position.

[0165] An aircraft as in any preceding clause, the sensor assembly comprising a first aft sensor and a second aft sensor each coupled to the aft bearing assembly.

[0166] An aircraft as in any preceding clause, the first aft sensor being coupled to the aft bearing assembly at a twelve o'clock position and the second aft sensor being coupled to the aft bearing assembly at a three o'clock position.

[0167] An aircraft as described in any preceding clause, wherein the sensor assembly includes one or more strain gauges or accelerometers.

[0168] An aircraft as described in any preceding clause, wherein the shaft is a low pressure turbine shaft.

[0169] An aircraft as claimed in any preceding clause, further comprising a load reduction device configured to fail when said propeller swirl is detected to be above said predetermined limit.

[0170] An aircraft as described in any preceding clause, further comprising means for controlling the pitch of each of said plurality of fan blades when said propeller swirl is detected to be above said predetermined limit.

[0171] A sensor assembly for detecting propeller swirl in a fan blade, the sensor assembly comprising a first front sensor, a second front sensor, a first rear sensor and a second rear sensor, each of the sensors being a strain gauge or an accelerometer.

[0172] A gas turbine engine having a sensor assembly according to any preceding clause.

[0173] A method of detecting propeller swirl using a sensor assembly according to the preceding clause.

[0174] A method of detecting propeller swirl of a fan blade of a gas turbine engine according to any preceding clause.

[0175] A method of detecting propeller swirl of a fan blade of an aircraft according to any preceding clause.

[0176] A method of detecting propeller swirl of a fan blade of an engine, the method comprising operating the engine, rotating a shaft in a circumferential direction during operation of the engine to rotate the fan blade, detecting a load on a bearing assembly supporting the shaft using a sensor, filtering the detected load to remove nominal and aerodynamic loads to generate a 1P load, comparing the 1P load to a predetermined limit, and detecting propeller swirl of the fan blade based on the comparison.

[0177] A method as described in the preceding clause, further comprising taking remedial action when propeller swirl is detected.

[0178] A method as in any preceding clause, filtering the detected load further comprising filtering using a bandpass filter.

[0179] A method as in any preceding clause, wherein the sensor is a strain gauge or an accelerometer.

[0180] A method as in any preceding clause, wherein the shaft is a low pressure turbine shaft.

[0181] A method as in any preceding clause, the engine being a non-ducted engine. A method as in any preceding clause, further comprising taking action when the IP load is above the predetermined limit.

[0182] A method as in any preceding clause, wherein the action is a passive action comprising failure of a load reduction device when the propeller swirl is above the predetermined limit.

[0183] A method as in any preceding clause, wherein the action is an active action comprising controlling the pitch of each of the plurality of fan blades when the propeller swirl is above the predetermined limit.

[0184] A method as in any preceding clause, wherein the detected load is a total load experienced by the mounted assembly, the total load being equal to the sum of a nominal engine operating load, an aerodynamic load and the IP load.

[0185] A method as in any preceding clause, wherein filtering the sensed load comprises subtracting each of the nominal engine operating load and the aerodynamic load from the total load to produce the IP load.

[0186] A method as in any preceding clause, wherein each of the total load, the nominal engine operating load, the aerodynamic load and the IP load comprises an axial force, a vertical force, a lateral force, a rolling moment, a pitch moment and a yaw moment.

[0187] A method as in any preceding clause, wherein the nominal engine loads include axial forces due to thrust, vertical forces due to gravity, pitch moments due to rotational motion, and yaw moments.

[0188] A method as defined in the preceding clause, wherein the lateral force and the rolling moment at said nominal engine load are both zero.

[0189] A method as in any preceding clause, wherein the aerodynamic loads include vertical and lateral forces each due to shear, a pitch moment and a yaw moment due to aerodynamics.

[0190] A method as defined in the preceding clause, wherein the aerodynamic load comprises zero axial force and zero rolling moment.

[0191] A method as in any preceding clause, wherein the 1P loads include vertical and lateral forces each due to shear, a pitch moment and a yaw moment due to 1P dynamic loads.

[0192] A method as defined in the preceding clause, wherein the aerodynamic load comprises zero axial force and zero rolling moment.

[0193] A method as in any preceding clause, wherein the aerodynamic load and the nominal load are determined from engine test data.

[0194] 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 for monitoring propeller swirl, characterized in that: The gas turbine engine comprises: a frame housing the gas turbine engine; a plurality of fan blades configured to rotate about a longitudinal centerline axis of the gas turbine engine; a mounting assembly coupling the frame to a structure; a sensor assembly coupled to the mounting assembly and having at least one sensor configured to detect a load on the mounting assembly, wherein the sensor assembly is configured to monitor propeller swirl in the plurality of fan blades by detecting the load on the mounting assembly; and feature configured to take corrective action when propeller swirl is detected above a predetermined limit.

2. The gas turbine engine according to claim 1, characterized in that in, The at least one sensor is a strain gauge or a load sensor.

3. The gas turbine engine according to claim 1, characterized in that in, The mounting assembly includes a front mounting member, a thrust mounting member link, and a rear mounting member, and the at least one sensor includes a first front sensor and a second front sensor on the front mounting member, a first rear sensor and a second rear sensor on the rear mounting member, and a thrust sensor on the thrust mounting member link.

4. The gas turbine engine according to claim 1, characterized in that in, The feature is a load reduction device configured to fail when the propeller swirl is detected to be above the predetermined limit.

5. The gas turbine engine according to claim 1, characterized in that in, The feature is a device configured to control the pitch of each of the plurality of fan blades when the propeller swirl is detected to be above the predetermined limit.

6. The gas turbine engine according to claim 1, characterized in that in, The frame includes a fan frame, and the mounting assembly includes a front mount coupling the fan frame to the structure, and wherein the at least one sensor is located on the front mount.

7. The gas turbine engine according to claim 6, characterized in that The front mount includes a first front mount link and a second front mount link, and the at least one sensor includes a first front sensor located on the first front mount link and a second front sensor located on the second front mount link.

8. The gas turbine engine according to claim 1, characterized in that in, The frame includes a turbine aft frame, and the mounting assembly includes an aft mount coupling the turbine aft frame to the structure, and wherein the at least one sensor includes a sensor located on the aft mount.

9. The gas turbine engine according to claim 8, characterized in that The rear mount includes a first rear mount link and a second rear mount link, and the at least one sensor includes a first rear sensor located on the first rear mount link and a second rear sensor located on the second rear mount link.

10. The gas turbine engine according to claim 8, characterized in that Further comprising a thrust mount link coupled to the rear mount, the at least one sensor comprising a thrust mount sensor located on the thrust mount link.

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