Sensor assembly for detecting propeller swirl in turbine engine

Abstract

The invention discloses a turbine engine. A turbine engine includes a plurality of fan blades configured to rotate about a longitudinal centerline axis of the turbine engine, a pitch actuator configured to control a pitch angle of each of the plurality of fan blades, a sensor assembly configured to detect a holding torque of the plurality of fan blades, and a controller configured to monitor propeller swirl in the plurality of fan blades by detecting the holding torque. The sensor assembly includes a first sensor coupled to the pitch actuator and configured to detect a holding torque of all of the plurality of fan blades, and a second sensor coupled to the pitch actuator and configured to detect a holding torque of all of the plurality of fan blades. The second sensor is coupled to the pitch actuator and is configured to detect a holding torque of at least one fan blade of the plurality of fan blades. The pitch actuator is configured to take a corrective action when it is detected that the propeller swirl is outside a predetermined limit.

Description

Technical Field

[0001] The present disclosure generally relates to a system for detecting propeller swirl in a turbine engine. Background Art

[0002] Turbine engines typically include a fan and a core section arranged in flow communication with each other. 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 It shows that according to the present disclosure Figure 1 An enlarged cross-sectional view of a portion of a gas turbine engine.

[0006] Figure 3A A fan blade (such as Figure 1 Schematic diagram of the forces on the fan blades of a turbine engine shown.

[0007] Figure 3B A fan blade (such as Figure 1 Schematic diagram of the forces on the fan blades of a turbine engine shown.

[0008] Figure 4 A method for detecting fan blades (such as Figure 1 Method of swirl of fan blades of a turbine engine as shown.

[0009] Figure 5 A graph of blade holding torque as a function of time is shown in accordance with the present disclosure.

[0010] Figure 6 The method for executing the present invention is shown Figure 4 An exemplary system of the method. DETAILED DESCRIPTION

[0011] 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.

[0012] 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.

[0013] As used herein, the terms "first," "second," "third," etc. may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0014] 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.

[0015] 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.

[0016] 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.

[0017] 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.

[0018] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0019] 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.

[0020] 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.

[0021] 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.

[0022] The present disclosure describes a detection engine (such as Figure 1 Systems and methods for swirl in fan blades of a gas turbine engine (described herein). Swirl is also referred to herein as prop whirl, 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 central 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 system that measures the holding torque on the fan blades to detect swirl in the fan blades. The sensor may include one or more accelerometers, one or more load sensors, one or more strain gauges, one or more pressure sensors, one or more conductive strips, or any combination thereof. The present disclosure provides a method for detecting swirl of a fan blade using sensor data.

[0023] 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. Figure 1The gas turbine engine 10 is a "three-stream engine" because its architecture provides three different thrust-generating airflow streams (labeled S1, S2, and S3) during operation, as further described below. Aspects of the present disclosure may be incorporated into any other suitable turbine engine (including, but not limited to, turbofan engines, propfan engines, turbojet engines, turboprop engines, and turboshaft engines, aviation-based turbine engines, marine turbine engines, land-based turbine engines, industrial turbine engines, power generation turbine engines, etc.).

[0024] 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.

[0025] 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.

[0026] 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.

[0027] 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.

[0028] 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 1In an embodiment, each fan blade 54 of the primary fan 52 may rotate about its respective central blade axis 57, for example in unison with each other. One or more actuators 58 are controlled to cause the fan blades 54 to pitch about their respective central blade axis 57. The pitch of the fan blades 54 is controlled based on the desired aerodynamics or engine performance. That is, one or more actuators 58 control the pitch of the fan blades 54 based on engine operating conditions, and may be controlled to change the pitch to achieve a predetermined engine performance. For example, as described herein, the fan blades 54 are controlled to change the aerodynamic force acting on the fan blades 54 and to reduce or prevent the propeller swirl acting thereon. In other embodiments, each fan blade 54 is fixed or cannot pitch about the central blade axis 57.

[0029] 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 their 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.

[0030] 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 1The 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.

[0031] 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.

[0032] 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 on the inner side of 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. In some cases, 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.

[0033] 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.

[0034] 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.

[0035] 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.

[0036] 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 .

[0037] 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.

[0038] 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.

[0039] Figure 1 The gas turbine engine 10 depicted in the figure is only used as 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 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)).

[0040] The controller 90 communicates with the turbine engine 10 for controlling various aspects of the turbine engine 10. For example, the controller 90 is in bidirectional communication with the turbine engine 10 for receiving signals from various sensors and control systems of the turbine engine 10, and for controlling components of the turbine engine 10, as described in further detail below. The controller 90 or components thereof may be located on the turbine engine 10, on the aircraft, or may be located remotely from each of the turbine engine 10 and the aircraft. The controller 90 may be a full authority digital engine control (FADEC) that controls various aspects of the turbine engine 10.

[0041] The controller 90 may be a stand-alone controller or may be part of an engine controller to operate various systems of the turbine engine 10. In this embodiment, the controller 90 is a computing device having one or more processors and a memory. The one or more processors may be any suitable processing device, including but not limited to a microprocessor, a microcontroller, an integrated circuit, a logic device, a programmable logic controller (PLC), an application specific integrated circuit (ASIC), or a field programmable gate array (FPGA). The memory may include one or more computer-readable media, including but not limited to non-transitory computer-readable media, computer-readable non-volatile media (e.g., flash memory), RAM, ROM, a hard drive, a flash drive, or other memory devices.

[0042] The memory may store information accessible by one or more processors, including computer-readable instructions executable by one or more processors. Instructions may be any set of instructions or sequence of instructions that cause one or more processors and the controller 90 to operate when executed by one or more processors. The controller 90, and more specifically, one or more processors, are programmed or constructed to perform these operations, such as the operations discussed further below. In some embodiments, instructions may be executed by one or more processors to cause one or more processors to perform any operations and functions that the controller 90 is constructed for, as will be further described below. Instructions may be software written in any suitable programming language or may be implemented in hardware. Additionally or alternatively, instructions may be executed in logically or virtually separate threads on a processor. The memory may further store data that may be accessed by one or more processors.

[0043] The technology discussed herein refers to computer-based systems, actions taken by computer-based systems, information sent to computer-based systems, and information sent from computer-based systems. Those of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for various possible configurations, combinations, and divisions of tasks and functions between and within components. For example, the processing discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and applications can be implemented on a single system, or can be distributed on multiple systems. Distributed components can operate sequentially or in parallel.

[0044] Figure 2 Shows Figure 1 The enlarged part is 100. Figure 1 As depicted, fan blades 54 (only one shown) are rotatably coupled to hub 53. Multiple fan blades 54 may be circumferentially spaced about hub 53 and longitudinal centerline axis 12. Thus, while one fan blade 54 is shown, more fan blades 54 may be provided. Fan blades 54 are coupled to fan shaft 56 ( Figure 1 ), the actuator 102 may be Figure 1 The actuator 102 can control the fan blade 54 relative to the longitudinal centerline axis 12 ( Figure 1 ) of the pitch angle. Controller 90( Figure 1 ) can control the actuator 102, thereby controlling the pitch angle of the fan blade 54. The actuator 102, also referred to herein as the pitch actuator 102, can control the feather position and the rough position of the fan blade 54. The actuator 102 includes a plurality of links to couple the blade pitch lever 110 of each fan blade 54 to the rotary linear actuator 112. The plurality of links may include a first link 104, a second link 106, and a third link 108. The blade pitch lever 110 includes a bearing assembly 114 to allow relative movement of the fan blade 54 relative to the static mounting members (e.g., the first static mounting member 116 and the second static mounting member 118).

[0045] The rotary linear actuator 112 is coupled to a plurality of links and includes a hydraulic cylinder 120 that defines a first chamber 122 and a second chamber 124. The first chamber 122 is separated from the second chamber 124 by a wall 126. During operation of the actuator 102, pressure may be supplied to or removed from the first chamber 122, the second chamber 124, or both to control the angle of the fan blade 54. A blade angle sensor 128 may be provided to measure the angle of the fan blade 54 relative to the longitudinal centerline axis 12 ( Figure 1). Controlling fan blades 54 using rotary linear actuators 112 is well known and will be understood by those of ordinary skill in the art that pressure applied to first chamber 122 (and / or removed from second chamber 124) causes first link 104, second link 106, and third link 108 to move axially in a first axial direction to rotate fan blades 54 in a first rotational direction. Similarly, pressure applied to second chamber 124 (and / or removed from first chamber 122) causes first link 104, second link 106, and third link 108 to move axially in a second axial direction (opposite to the first axial direction) to rotate fan blades 54 in a second rotational direction opposite to the first rotational direction.

[0046] Brief reference Figure 3A and Figure 3B During operation, fan blades 54 rotate in a clockwise direction 308 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 3A and Figure 3B 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 300 to the three o'clock position 302, to the six o'clock position 304, to the nine o'clock position 306, and back to the twelve o'clock position 300. 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, in the axial direction 310 and the vertical direction 312 (e.g., relative to the ground or horizontal reference), 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).

[0047] Thus, during operation at the 1P frequency (eg, the base RPM of the fan blade 54), the rotating fan blade 54 is subject to two phenomena depending on the circumferential position of the blade 54, such as Figure 3A and Figure 3B As shown. Figure 3A , fan blade 54 experiences shear and a downward moment in vertical direction 350, which causes fan blade 54 to yaw rightward in yaw direction 360 (e.g., twist or oscillate about vertical direction 350). The vertically directed shear and rightward directed yaw cause fan blade 54 to pitch upward. Figure 3B , fan blades 54 also experience shear and lateral movement in a lateral direction 370, which causes fan blades 54 to move in a pitch direction 380 (e.g., in a direction parallel to the pitch direction 380). Figure 3AThe fan blade 54 is pitched downwardly in a direction opposite to the vertical direction 350 of the fan blade 54. The rightward directed shear and downward directed pitch causes 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 frequency 1P excites the backward swirl motion 390 in the fan blade 54. Since the fan blade 54 is constrained at the base of the fan blade 54 (e.g., as shown in FIG. 1 ), the fan blade 54 is tilted downwardly in a direction opposite to the vertical direction 350 of the fan blade 54. The rightward directed shear and downward directed pitch causes 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 frequency 1P excites the backward swirl motion 390 in the fan blade 54. Figure 2 ), so as previously described, when the structure holding the fan blade 54 in place acts to maintain the fan blade 54 at a predetermined pitch angle, the force pulling the blade that causes the propeller swirl also causes a change in the holding torque of the fan blade 54. Therefore, the sensor assembly disclosed herein allows the propeller swirl to be monitored by measuring the amount of torque required to hold the base of the fan blade 54 at a predetermined pitch angle and any changes in this measured torque.

[0048] Return reference, Figure 2 A sensor assembly 200 and a second static mounting member 118 associated with the actuator 102 are shown. The sensor assembly 200 may include one or more sensors. The one or more sensors may include a pressure sensor 202 coupled to the rotary linear actuator 112 and associated with the first chamber 122, a pressure sensor 204 coupled to the rotary linear actuator 112 and associated with the second chamber 124, a strain gauge 206 (also referred to herein as a strain sensor) coupled to the blade pitch rod 110, a conductive tape 208 coupled between the second link 106 and the blade pitch rod 110, a load sensor 210 coupled to the bearing assembly 114, an accelerometer 212 coupled to the second static mounting member 118, and an accelerometer 214 mounted on the third link 108 coupling the second link 106 to the blade pitch rod 110.

[0049] Although the sensor assembly 200 is shown and described as including the above-described sensors, any combination or sub-combination of sensors may be present, and / or there may be fewer sensors than those shown and described. For example, the sensor assembly 200 may include only one or two pressure sensors (e.g., pressure sensor 202 may be the first pressure sensor 202 and / or pressure sensor 204 may be the second pressure sensor 204). In another example, the sensor assembly 200 includes only one or two pressure sensors and the strain gauge 206. In another example, the sensor assembly 200 includes only one or two accelerometers (e.g., accelerometer 212 may be the first accelerometer 212 and / or accelerometer 214 may be the second accelerometer 214). The above examples are not exhaustive, and any single sensor or combination of sensors described herein may form the sensor assembly 200.

[0050] The sensors of sensor assembly 200 detect various parameters indicative of blade holding torque. Blade holding torque is the torque required to maintain a desired pitch angle of fan blade 54. As will be discussed in greater detail, the blade holding torque sensed by sensor assembly 200 is related to the swirl pattern of fan blade 54. That is, the swirl pattern causes changes in the aerodynamic forces on fan blade 54, which results in changes in blade holding torque and changes in bearing loads.

[0051] refer to Figure 2 , the sensor assembly 200 may include a first pressure sensor 202 and a second pressure sensor 204 to measure the pressure change in the hydraulic cylinder 120 required to provide the holding torque for the fan blades 54. The first pressure sensor 202 and the second pressure sensor 204 are located in the hydraulic cylinder 120 to measure the pressure within the hydraulic cylinder 120. The first pressure sensor 202 is located in the first chamber 122 to detect the pressure change in the first chamber 122. The second pressure sensor 204 is located in the second chamber 124 to detect the pressure change in the second chamber 124. The rotary linear actuator 112 includes a single hydraulic cylinder 120, and therefore, the first pressure sensor 202 and the second pressure sensor 204 provide data related to all fan blades 54 combined. The pressure change measured in the hydraulic cylinder 120 provides the holding torque required for all fan blades 54 combined. The first pressure sensor 202 and the second pressure sensor 204 are assembled to the hydraulic cylinder 120 to measure the change in cylinder pressure, which provides insight into the change in the pure torsional load on all fan blades 54 combined. First pressure sensor 202, second pressure sensor 204, or both may be the only sensors in sensor assembly 200, or may be provided in conjunction with any one or more of the sensors described herein. Although two pressure sensors are depicted, more or fewer pressure sensors may be provided.

[0052] Hydraulic cylinder 120 provides actuation for actuator 102 to change the pitch of fan blades 54. When set to a predetermined pitch angle, the pressure change within hydraulic cylinder 120 indicates that the holding torque required to maintain fan blades 54 at the pitch angle has changed. Therefore, the measurement of the pressure change within one or two chambers (e.g., first chamber 122 and / or second chamber 124) of hydraulic cylinder 120 is an indication of the holding torque required to maintain actuator 102 and fan blades 54 at the desired pitch angle. Therefore, first pressure sensor 202 and second pressure sensor 204 measure the holding torque required to maintain all fan blades 54 at the desired pitch angle.

[0053] Continue to refer Figure 2, the sensor assembly 200 may include a strain gauge 206 on each blade pitch rod 110. The strain gauge 206 may be located on the circumferential surface of the blade pitch rod 110. The strain gauge 206 measures the strain on each blade pitch rod 110 (e.g., at the root of each blade of the fan blade 54). Therefore, each strain gauge 206 provides information related to the corresponding fan blade 54. Placing a strain gauge 206 on each blade pitch rod 110 allows for personalized measurement of the holding torque of each fan blade 54. The strain gauge 206 may be a wireless strain gauge. The strain gauge 206 provides strain measurement in the torsional direction, which provides a torque measurement in the corresponding fan blade 54. The strain gauge 206 is selected from a strain gauge that provides reliable measurement of the rotating component. In some examples, redundancy can be added to the sensor assembly 200 by including three sensors one hundred and twenty degrees apart on the same blade pitch rod 110 or by including two sensors ninety degrees apart. Strain gauge 206 may be the only sensor in sensor assembly 200, or may be provided in combination with any one or more of the sensors described herein. Although a single strain gauge 206 on each blade pitch rod 110 is described, more or fewer strain gauges may be provided.

[0054] Blade pitch lever 110 controls the pitch angle of corresponding fan blade 54 via actuator 102 (as previously described, each fan blade 54 has a pitch lever 110). Thus, the measurement of strain in the torsional direction indicates the holding torque that maintains fan blade 54 at a predetermined pitch angle.

[0055] The sensor assembly 200 may include a conductive strip 208 between the second link 106 and the blade pitch rod 110. The conductive strip 208 measures the force change in the third link 108 connected to the blade pitch rod 110. The conductive strip 208 is therefore disposed at each fan blade 54 to provide information related to the corresponding fan blade 54. The force change measured by the conductive strip 208 is related to the holding torque of each fan blade 54. The conductive strip 208 may be the only sensor in the sensor assembly 200, or may be provided in combination with any one or more sensors described herein. Although a single conductive strip 208 coupled to each blade pitch rod 110 is described, more or fewer conductive strips may be provided.

[0056] During pitching of fan blade 54, first link 104, second link 106, and third link 108 may be axially moved relative to blade pitch rod 110 by actuator 102. That is, links 104, 106, and 108 are moved axially forward and axially backward to control the pitching of fan blade 54 between various feathered positions and coarse positions. Conductive tape 208 may thus be placed in compression or tension during movement of second link 106 relative to blade pitch rod 110. When the stress (e.g., compression or tension) in conductive tape 208 changes, the resistance changes and can be detected. Thus, the change in resistance indicates a change in holding torque.

[0057] The sensor assembly 200 may include a load sensor 210 at each bearing assembly 114. The load sensor 210 is located below the outer raceway of the bearing in the bearing assembly 114 (e.g., on a surface opposite to the raceway surface that contacts the bearing). The load sensor 210 measures the radial load and axial load at the outer raceway of the bearing assembly 114, which provides an indication of aerodynamic changes on each fan blade 54. The load sensor 210 measures the load (radial and axial) at the blade root (e.g., the portion of the blade assembly that couples the fan blade 54 to the blade pitch rod 110). The load sensor 210 may be located at each blade root of the fan blade 54. Thus, data is collected at all blade positions to provide an indication of the swirl pattern at each fan blade 54. The load sensor 210 may be the only sensor in the sensor assembly 200, or may be provided in combination with any one or more sensors described herein. Although a single load sensor 210 coupled to each bearing assembly 114 is described, more or fewer load sensors may be provided.

[0058] As fan blade 54 rotates about longitudinal centerline axis 12, the load applied by fan blade 54 on bearing assembly 114 changes. This change in aerodynamic load results in a change in the bearing load on bearing assembly 114. Load sensor 210 measures the change in bearing load by measuring the radial pull on fan blade 54, which is an indication of the holding torque on fan blade 54.

[0059] The sensor assembly 200 may include a first accelerometer 212 and a second accelerometer 214. The first accelerometer 212 and the second accelerometer 214 measure the acceleration of the component in which they are located in a specific predetermined direction. The second accelerometer 214 is located on the third link 108, and the first accelerometer 212 is located on a non-rotating component (such as the second static mounting member 118) near the outlet guide vane. The second accelerometer 214 located on the third link 108 detects pure torsional vibrations in each fan blade 54. The first accelerometer 212, the second accelerometer 214, or both the first accelerometer 212 and the second accelerometer 214 may be the only sensors in the sensor assembly 200, or may be provided in combination with any one or more sensors described herein. Although two accelerometers are depicted, more or fewer accelerometers may be provided.

[0060] The second accelerometer 214 located on the rotating component measures the frequency of the rotating system. The first accelerometer 212 located on the static component can measure other ambient frequencies generated by the engine 10 and is used to filter noise from the frequency signal obtained from the second accelerometer 214.

[0061] Thus, the sensors of sensor assembly 200 may provide data for all fan blades 54 combined (e.g., pressure sensors), or may provide data for each individual fan blade 54 (e.g., strain gauges, accelerometers, load sensors, and conductive strips). Sensor assembly 200 may combine one or more sensors to detect multiple parameters, thereby improving or enhancing gas turbine engine 10 (e.g., using a single sensor). Figure 1 ) in the detection of swirl patterns in the gas turbine engine 10. In some examples, sensors that detect parameters of the combined fan blades 54 can be combined with sensors that detect parameters of the individual fan blades 54 to detect Figure 1 ) in a swirl pattern. For example, sensor assembly 200 may include one or more pressure sensors (e.g., one or both of pressure sensor 202 and pressure sensor 204) and strain gauge 206. In this example, other sensors described herein may be optional and may be omitted. This is only one example, and other combinations of sensors are contemplated. The sensor assembly 200 of the present disclosure may include radio frequency identification (RFID), radio waves, short-range wireless technologies (e.g., ), near field communication (NFC) or other types of wireless communication to transmit data from sensors located on rotating components.

[0062] The sensor system 200 and a controller (such as Figure 1Controller 90 communicates with controller 90 of sensor system 200 to provide detection and monitoring of the presence of propeller swirl and, in some cases, provide control of the engine to mitigate or reduce propeller swirl. For example, controller 90 is configured to receive information and data from sensors of sensor system 200. Controller 90 collects data and analyzes the data, as discussed in more detail below, to detect and monitor the presence of propeller swirl. Controller 90 may also take action (e.g., as described with respect to Figure 4 The corrective action described in step 412 of the described method 400).

[0063] When fan blade 54( Figure 1 ) rotates about the longitudinal centerline axis 12, the mass imbalance created by the rotation of the fan blades 54 may cause propeller swirl. When propeller swirl occurs, the fan blades 54 may not rotate precisely along the longitudinal centerline axis 12 and / or the angle of attack of the fan blades may change. The change in angle of attack will result in a change or variation in the holding torque required to maintain the desired pitch angle of the fan blades 54. The sensor assembly 200 measures this holding torque at a predetermined pitch angle (e.g., set by a pitch actuator), which can then be used to control the fan blades 54 via the controller 90 ( Figure 1 ) determines the presence of propeller swirl. That is, if the systems and methods herein detect that the change in the holding torque of the blades is outside of an acceptable range, then the systems and methods have detected the presence of propeller swirl.

[0064] As previously described, sensor assembly 200 (eg, pressure sensor, strain sensor, and / or conductive tape) measures the total holding torque ht of each fan blade 54. total . Total holding torque ht total It is limited by the relation (1):

[0065] ht total = ht nominal + ht 1P (1)

[0066] In equation (1), ht nominal It represents the nominal holding torque determined by the engine test rig or laboratory test, ht 1P is the holding torque due to the propeller swirl and loads generated by the 1P frequency (e.g., due to the load at the base RPM of the blades 54). The nominal holding torque is the torque required to hold the blades in position during normal engine operation in all phases of flight without any abnormal behavior (e.g., vibration levels exceeding predetermined limits set by the engine operator). The total holding torque ht total Allows determination of the holding torque ht 1P , as shown in equation (2):

[0067] ht 1P = ht total -ht nominal (2)

[0068] Likewise, the sensor assembly 200 (eg, a load cell) measures the total bearing load Rad of each fan blade 54. total The total radial load Rad on the outer raceway of the bearing total It is limited by the relation (3):

[0069] Rad total = Rad nominal + Rad 1P (3)

[0070] In equation (3), Rad nominal Represents the normal radial load experienced at the bearing during normal operation and is determined by test bench or laboratory testing of the engine, Rad 1P Radial load due to propeller swirl and 1P load. total Allows determination of the bearing load Rad 1P , as shown in equation (4):

[0071] Rad 1P = Rad total - Rad nominal (4)

[0072] refer to Figure 4 , controller 90( Figure 1) is programmed to perform method 400 to detect a swirl pattern of a fan blade 54 in a gas turbine engine 10 including a sensor assembly 200, and to take corrective action when a swirl pattern is detected. At step 402, the fan blades 54 are placed in operation (e.g., rotation begins), and aerodynamic forces act on each fan blade 54 as previously described. At step 404, sensors of the sensor assembly 200 detect and record data. The sensors may record pure torsional vibrations along the pitch axis of the fan blades and / or varying loads at bearing locations, as described herein. This may be accomplished by any combination of sensors described herein. The sensors detect and record this data at a 1P frequency (e.g., the base RPM of the blades 54). The sensors transmit the data to a controller, where the data may be analyzed and stored. At step 406, the controller is used to compare the measured data with baseline data 408 obtained from a test bench test. This comparison may result in detection of a propeller swirl pattern at step 410. Detecting a propeller swirl mode at step 410 may (1) indicate instability if the signal amplitude continues to increase beyond a predetermined acceptable range, (2) indicate the presence of a propeller swirl mode if the signal is outside of a predetermined acceptable range, or (3) indicate the absence of instability if the signal amplitude is within a predetermined acceptable range.

[0073] After the propeller swirl pattern is detected at step 410, and if appropriate, the controller may initiate corrective action or remedial action to reduce or eliminate the propeller swirl at step 412. For example, if the propeller swirl pattern is detected as being outside of a predetermined range or predetermined limit, corrective action may be taken at step 412. If the propeller swirl pattern is detected as being an acceptable value (e.g., within an acceptable predetermined range or below an acceptable predetermined limit) at step 410, no action is required.

[0074] An exemplary corrective action taken at step 412 may be to utilize pitch actuator 102 ( Figure 2 ) for fan blades (e.g., Figure 1 Pitch control of the fan blades 54 described above. Pitch control can be micro-pitch of the fan blades at a specific frequency to offset the changes in aerodynamic force caused by the propeller swirl. Micro-pitch is achieved by providing a pulsating blade holding torque to the blades. In an example, the micro-pitch of each blade can be independently controlled relative to each of the remaining blades. In an example, all blades are controlled, and each blade is controlled independently of the remaining blades.

[0075] Micro pitch refers to changing the pitch angle of the blades. Micro pitch can suppress propeller swirl by changing the blade pitch angle to obtain a rotational load vector on the blade, thereby alleviating or reducing the propeller swirl. The amount of change in the pitch angle (e.g., the micro pitch amount) depends on the size of the rotational load vector to be generated to suppress the propeller swirl. The amount can be determined by a controller. In some examples, micro pitch of the angle includes changing the pitch angle by less than 0.5 degrees. The frequency of the micro pitch is at the 1P frequency. Micro pitch is performed until the propeller swirl is alleviated (e.g., the value is within a predetermined acceptable range or below a predetermined threshold level), provided that the vibration is not increased from the existing level due to the micro pitch. Therefore, the propeller swirl continues to be monitored during micro pitch, so that the micro pitch can be changed to further alleviate the propeller swirl, or stopped when the propeller swirl is no longer detected. Monitoring includes repeatedly detecting the holding torque of all multiple fan blades and each individual fan blade, and determining whether the holding torque indicates a propeller swirl. Monitoring is performed by a controller (e.g., Figure 1 Controller 90) is used.

[0076] As mentioned earlier, the propeller swirl produces changes in the aerodynamic forces on the blades, which also result in changes in the blade holding torque. Figure 5 , graph 500 depicts blade holding torque as a function of time. It can be seen that there is a nominal torque that represents the required blade holding torque. During engine operation, blade holding torque 502 can deviate from the nominal torque. It can be seen that due to changing aerodynamic forces, blade holding torque 502 also changes. The changes shown in graph 500 represent the changing aerodynamic forces associated with the propeller swirl.

[0077] To correct for propeller swirl, the fan blades may be micro-pitch. Micro-pitch includes slightly changing (e.g., less than 0.5 degrees) the pitch angle of each fan blade independently of the other fan blades at a predetermined frequency and a predetermined phase in order to offset the changing aerodynamic forces acting on the fan blades due to the presence of propeller swirl. The pitch angle is the angle of the fan blade relative to the longitudinal pitch axis of the fan blade. The predetermined frequency and predetermined phase may be set to values ​​associated with the 1P load, and / or may be determined by bench testing.

[0078] Micro pitch can also be used to create rotational load vectors (radial load vectors and / or thrust load vectors) on the fan blades so that the rotational load vectors offset smaller rotational imbalance loads in the fan blades, which can reduce propeller swirl by forcing the loads closer to the theoretical center of rotation.

[0079] Figure 6 A method for utilizing the controller 90 ( Figure 1) is an exemplary system for implementing method 400. Controller 90 may implement steps 406, 410, and 412 of method 400. The system includes a general computing device 600, which includes a processing unit (CPU or processor) 620 and a system bus 610, which connects various system components including system memory 630 (such as read-only memory (ROM) 640 and random access memory (RAM) 650) to the processor 620. The system may include a cache of high-speed memory directly connected to the processor 620, adjacent to the processor 620, or integrated as a part of the processor 620. The system copies data from the system memory 630 and / or storage device 660 to the cache for rapid access by the processor 620. In this way, the cache provides a performance improvement that avoids delays in the processor 620 while waiting for data. These modules and other modules may control or be configured to control the processor 620 to perform various actions. Other system memories 630 may also be used. The system memory 630 may include a variety of different types of memory with different performance characteristics. The present disclosure may operate on a computing device 600 having more than one processor 620, or on a group or cluster of computing devices networked together to provide greater processing capabilities. The processor 620 may include any general purpose processor and hardware modules or software modules (such as module 1 662, module 2 664, and module 3 666 stored in the storage device 660 and configured to control the processor 620), as well as a dedicated processor where software instructions are incorporated into the actual processor design. The processor 620 may essentially be a completely independent computing system, including multiple cores or processors, buses, memory controllers, caches, etc.

[0080] The system bus 610 can be any of several types of bus structures, including a memory bus or memory controller, a peripheral bus, and a local bus using any of a variety of bus architectures. A basic input / output (BIOS) stored in a ROM 640 or the like can provide basic routines that facilitate the transfer of information between elements within the computing device 600, such as during startup. The computing device 600 also includes a storage device 660, such as a hard disk drive, a magnetic disk drive, an optical disk drive, a tape drive, etc. The storage device 660 may include software modules 662, 664, 666 for controlling the processor 620. Other hardware or software modules may be envisioned. The storage device 660 is connected to the system bus 610 via a drive interface. The drive and associated computer-readable storage media provide non-volatile storage of computer-readable instructions, data structures, program modules, and other data for the computing device 600. In one aspect, a hardware module that performs a specific function includes a software component stored in a tangible computer-readable storage medium, which is combined with a hardware component (such as a processor 620, a system bus 610, a display 670, etc.) required to perform the function. In another aspect, the system may use a processor and a computer-readable storage medium to store instructions that, when executed by a processor (e.g., one or more processors), cause the processor to perform a method or other specific actions. Depending on the type of device (such as whether the computing device 600 is a small handheld computing device, a desktop computer, or a computer server), basic components and appropriate variations are contemplated.

[0081] Although the exemplary embodiments described herein employ a hard disk or storage device 660, other types of computer-readable media that can store data accessible by a computer (such as magnetic cassettes, flash memory cards, digital versatile disks, cassettes, random access memory (RAM) 650, and read-only memory (ROM) 640) may also be used in the exemplary operating environment. Tangible computer-readable storage media, computer-readable storage devices, or computer-readable memory devices expressly exclude media (such as transient waves, energy, carrier signals, electromagnetic waves) and signals themselves.

[0082] To enable user interaction with computing device 600, input device 690 represents any number of input mechanisms, such as a microphone for voice, a touch-sensitive screen for gesture or graphic input, a keyboard, a mouse, motion input, voice, etc. Output device or display 670 may also be one or more of a plurality of output mechanisms known to those skilled in the art. In some cases, a multimodal system enables a user to provide multiple types of input to communicate with computing device 600. Communication interface 680 typically governs and manages user input and system output. There is no restriction on operating on any particular hardware arrangement, and therefore, the basic features here can be easily replaced by improved hardware or firmware arrangements when developed.

[0083] The technology discussed herein refers to computer-based systems, actions taken by computer-based systems, and information sent to and from computer-based systems. Those of ordinary skill in the art will recognize that the inherent flexibility of computer-based systems allows for various possible configurations, combinations, and divisions of tasks and functions between and among components. For example, the processing discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and applications can be implemented on a single system or distributed on multiple systems. Distributed components can operate sequentially or in parallel.

[0084] Therefore, the sensor assembly of the present disclosure detects propeller swirl. As discussed, the blade holding torque is directly related to the swirl pattern. That is, the propeller swirl pattern causes a change in the aerodynamic force on the fan blade 54, which, as previously described, causes a change in the blade holding torque and a change in the bearing load. Therefore, the measurement of the blade holding torque (provided by the sensor assembly disclosed herein) provides detection of the propeller swirl pattern. In other words, the sensor assembly 200 of the present disclosure detects propeller swirl by detecting changes in the blade holding torque (the torque required to maintain the desired pitch angle of the fan blade). By placing the sensor in a position capable of detecting the blade holding torque, the sensed data can be associated to detect propeller swirl, because the change in the holding torque at the 1P frequency (e.g., the basic rpm of the blade) is a direct measurement of the change in the aerodynamic force on the blade. The detection described herein can be performed at the test bench test level to predict normal operation with field instability or imbalance.

[0085] Further aspects are provided by the subject matter of the following clauses.

[0086] A turbine engine, comprising: a plurality of fan blades, the plurality of fan blades being configured to rotate about a longitudinal centerline axis of the turbine engine; a pitch actuator, the pitch actuator being configured to control the pitch angle of each of the plurality of fan blades; a sensor assembly, the sensor assembly being configured to detect a holding torque of the plurality of fan blades, wherein the holding torque represents the amount of torque required to maintain each of the plurality of fan blades at a predetermined pitch angle; and a controller, the controller being configured to monitor propeller swirl in the plurality of fan blades by detecting the holding torque. The sensor assembly comprises: a first sensor, the first sensor being coupled to the pitch actuator and being configured to detect the holding torque of all of the plurality of fan blades; and a second sensor, the second sensor being coupled to the pitch actuator and being configured to detect the holding torque of at least one of the plurality of fan blades. The pitch actuator is configured to take corrective action when propeller swirl is detected to be outside of a predetermined limit.

[0087] A turbine engine according to the preceding clause, wherein the first sensor is a pressure sensor and the second sensor is one or more of a strain gauge, a conductive tape, a load sensor or an accelerometer.

[0088] A turbine engine according to any preceding clause, further comprising a third sensor, the third sensor being one or more of a conductive tape, a load sensor or an accelerometer.

[0089] A turbine engine as described in any preceding clause, wherein said pitch actuator is configured to micro-pitch said plurality of fan blades as said corrective action.

[0090] A turbine engine as claimed in any preceding clause, wherein the pitch actuator comprises a hydraulic cylinder and the first sensor is a pressure sensor configured to detect a pressure change in the hydraulic cylinder of the pitch actuator.

[0091] A turbine engine according to any preceding clause, wherein the pressure sensors are first and second pressure sensors, and wherein the first pressure sensor is coupled to a first chamber of the hydraulic cylinder and the second pressure sensor is coupled to a second chamber of the hydraulic cylinder.

[0092] A turbine engine as described in any preceding clause, wherein the second sensor is a strain gauge.

[0093] A turbine engine as described in any preceding clause, wherein the strain gauge is coupled to a blade pitch lever coupled to the at least one fan blade.

[0094] According to the turbine engine described in any preceding item, the strain gauge is a plurality of strain gauges, each strain gauge is connected to a corresponding blade pitch rod, the corresponding blade pitch rod is connected to a corresponding fan blade among the plurality of fan blades, and each strain gauge is constructed to detect strain in the corresponding fan blade.

[0095] A turbine engine as described in any preceding clause, wherein the second sensor is a conductive strip.

[0096] The turbine engine according to any preceding clause, the pitch actuator further comprising a connecting rod and a blade pitch lever coupled to each fan blade of the plurality of fan blades, the conductive strip being coupled between at least one blade pitch lever and at least one connecting rod of the corresponding fan blade.

[0097] According to the turbine engine described in any of the preceding items, the pitch actuator further includes a connecting rod and a blade pitch rod connected to each of the plurality of fan blades, the conductive strip is a plurality of conductive strips, each conductive strip is connected between the blade pitch rod and the connecting rod of a corresponding fan blade, and each conductive strip is constructed to detect a change in resistance associated with the corresponding fan blade.

[0098] A turbine engine as described in any preceding clause, wherein the second sensor is a load sensor.

[0099] A turbine engine as recited in any preceding clause, wherein the load sensor is a plurality of load sensors, each load sensor coupled to a respective one of the plurality of fan blades, and each load sensor configured to detect a change in a bearing load of the respective fan blade.

[0100] A turbine engine as described in any preceding clause, wherein the load sensor is coupled to a root of the at least one fan blade.

[0101] A turbine engine as described in any preceding clause, wherein the load sensor is coupled to an outer race of a bearing assembly of the fan blade.

[0102] A turbine engine as described in any preceding clause, wherein the second sensor is an accelerometer.

[0103] A turbine engine as described in any preceding clause, the accelerometer being coupled to a rotating component in the turbine engine.

[0104] A turbine engine as claimed in any preceding clause, wherein the rotating component is a connecting rod of the pitch actuator.

[0105] A turbine engine according to any preceding clause, wherein the accelerometer is a plurality of accelerometers, each accelerometer being coupled to a respective rotating component coupled to a respective fan blade of the plurality of fan blades, and each accelerometer being configured to detect a frequency of the respective rotating component.

[0106] A turbine engine according to any preceding clause, wherein the accelerometer is a first accelerometer, the turbine engine further comprising a second accelerometer coupled to a static component of the turbine engine, the second accelerometer being configured to detect ambient frequencies of the turbine engine to be filtered from a signal generated by the first accelerometer.

[0107] A turbine engine as described in any preceding clause, wherein the first sensor is a pressure sensor and the second sensor is one or more of a strain gauge, a conductive tape, a load sensor, or an accelerometer.

[0108] A turbine engine as described in any preceding clause, further comprising a third sensor, the third sensor being one or more of a conductive tape, a load sensor, or an accelerometer, and the second sensor being a strain gauge.

[0109] A turbine engine as described in any preceding clause, wherein the first sensor is a pressure sensor.

[0110] A turbine engine as described in any preceding clause, further comprising a third sensor, the third sensor being one or more of a strain gauge, a load sensor, or an accelerometer, and the second sensor being a conductive tape.

[0111] A turbine engine as described in any preceding clause, further comprising a third sensor, the third sensor being one or more of a strain gauge, a conductive tape, or an accelerometer, and the second sensor being a load sensor.

[0112] A turbine engine as described in any preceding clause, further comprising a third sensor, the third sensor being one or more of a strain gauge, a load sensor, or a conductive tape, and the second sensor being an accelerometer.

[0113] A turbine engine as described in any preceding clause, further comprising a conductive tape, a load sensor and an accelerometer, and wherein the second sensor is a strain gauge.

[0114] A turbine engine as described in any preceding clause, wherein said pitch actuator is configured to micro-pitch said plurality of fan blades as said corrective action.

[0115] A sensor assembly for detecting propeller swirl in a fan blade, the sensor assembly comprising one or more of a pressure sensor configured to detect a holding torque of the fan blade, a strain gauge configured to detect the holding torque of the fan blade, a conductive strip configured to detect the holding torque of the fan blade, an accelerometer configured to detect the holding torque of the fan blade, or a load sensor configured to detect the holding torque of the fan blade, the pressure sensor, the strain gauge, the conductive strip, the load sensor, and the accelerometer each being coupled to a pitch actuator of the fan blade.

[0116] A sensor assembly for detecting propeller swirl in a fan blade, the sensor assembly comprising two or more of a pressure sensor configured to detect a holding torque of the fan blade, a strain gauge configured to detect the holding torque of the fan blade, a conductive strip configured to detect the holding torque of the fan blade, an accelerometer configured to detect the holding torque of the fan blade, or a load sensor configured to detect the holding torque of the fan blade, the pressure sensor, the strain gauge, the conductive strip, the load sensor, and the accelerometer each being coupled to a pitch actuator of the fan blade.

[0117] A method for detecting propeller swirl of a fan blade, the method comprising controlling the fan blade to a predetermined pitch angle, rotating the fan blade in a circumferential direction, detecting a holding torque required to maintain the fan blade in the circumferential direction using a sensor, comparing the holding torque with baseline data associated with the fan blade, detecting propeller swirl of the fan blade based on the comparison, and taking corrective action to mitigate the propeller swirl.

[0118] According to the method of the preceding clause, propeller swirl is detected when the holding torque is outside a predetermined range.

[0119] A method as in any preceding clause, wherein no propeller swirl is detected when the holding torque is within a predetermined range.

[0120] A method as in any preceding clause, detecting the holding torque comprising sensing a pressure change in a pitch actuator configured to provide the predetermined pitch angle.

[0121] A method as in any preceding clause, detecting the holding torque comprising sensing strain in a pitch actuator configured to provide the predetermined pitch angle.

[0122] A method as in any preceding clause, detecting the holding torque comprising sensing a change in a conduction band of a pitch actuator configured to provide the predetermined pitch angle.

[0123] A method as in any preceding clause, detecting the holding torque comprising sensing a frequency in a rotating component of a pitch actuator configured to provide the predetermined pitch angle.

[0124] A method as in any preceding clause, further comprising removing noise from the sensing generated signal.

[0125] A method as in any preceding clause, wherein the fan blade is a plurality of fan blades.

[0126] The method of the preceding clause, wherein the sensing comprises individually sensing the holding torque of each of the fan blades and sensing the holding torque of all of the plurality of fan blades together.

[0127] A method as in any preceding clause, detecting the holding torque comprises sensing a total holding torque with the sensor, and determining the holding torque due to propeller swirl by subtracting a nominal holding torque from the total holding torque.

[0128] A method as in any preceding clause, further comprising measuring a total bearing load on a bearing assembly supporting the fan blade, and determining a load on the bearing assembly due to propeller swirl by subtracting a normal bearing load from the total bearing load.

[0129] A method as in any preceding clause, further comprising micro-pitching the fan blades to counteract the holding torque associated with propeller swirl.

[0130] A method of detecting propeller swirl in an engine according to any preceding clause.

[0131] A method of detecting propeller swirl using a sensor assembly according to any preceding clause.

[0132] 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 turbine engine, characterized in that: include: a plurality of fan blades configured to rotate about a longitudinal centerline axis of the turbine engine; a pitch actuator configured to control a pitch angle of each of the plurality of fan blades; a sensor assembly configured to detect a holding torque of the plurality of fan blades, wherein the holding torque represents an amount of torque required to maintain each of the plurality of fan blades at a predetermined pitch angle, the sensor assembly comprising: a first sensor coupled to the pitch actuator and configured to detect the holding torque of all of the plurality of fan blades; and a second sensor coupled to the pitch actuator and configured to detect the holding torque of at least one fan blade of the plurality of fan blades; and a controller configured to monitor propeller swirl in the plurality of fan blades by detecting the holding torque, Wherein the pitch actuator is configured to take corrective action when propeller swirl is detected to be outside predetermined limits.

2. The turbine engine according to claim 1, characterized in that in, The first sensor is a pressure sensor and the second sensor is one or more of a strain gauge, a conductive tape, a load sensor, or an accelerometer.

3. The turbine engine according to claim 1, characterized in that: in, The pitch actuator is configured to fine-pitch the plurality of fan blades as the corrective action.

4. The turbine engine according to claim 1, characterized in that: in, The pitch actuator includes a hydraulic cylinder, and the first sensor is a pressure sensor configured to detect a pressure change in the hydraulic cylinder of the pitch actuator.

5. The turbine engine according to claim 4, characterized in that in, The pressure sensors are first and second pressure sensors, and wherein the first pressure sensor is coupled to a first chamber of the hydraulic cylinder and the second pressure sensor is coupled to a second chamber of the hydraulic cylinder.

6. The turbine engine according to claim 1, characterized in that in, The second sensor is a strain gauge.

7. The turbine engine according to claim 6, characterized in that in, The strain gauge is coupled to a blade pitch lever coupled to the at least one fan blade.

8. The turbine engine according to claim 7, characterized in that in, The strain gauge is a plurality of strain gauges, each strain gauge is coupled to a respective blade pitch lever coupled to a respective fan blade of the plurality of fan blades, and wherein each strain gauge is configured to detect strain in the respective fan blade.

9. The turbine engine according to claim 1, characterized in that in, The second sensor is a conductive strip.

10. The turbine engine according to claim 9, characterized in that The pitch actuator further includes a connecting rod and a blade pitch lever coupled to each fan blade of the plurality of fan blades, wherein the conductive strap is coupled between at least one blade pitch lever and at least one connecting rod of the corresponding fan blade.

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