Patent Application
20250230753
The present disclosure relates generally to fan actuation systems for turbofan engines.
Turbofan engines, for example, for an aircraft, generally include a fan having fan blades and a turbo-engine arranged in flow communication with one another. Some turbofan engines include a fan actuation system for actuating the fan blades of the fan.
The foregoing and other features and advantages will be apparent from the following, more particular, description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein like reference numbers generally indicate identical, functionally similar, or structurally similar elements.
Features, advantages, and embodiments of the present disclosure are set forth or apparent from a consideration of the following detailed description, drawings, and claims. Moreover, the following detailed description is exemplary and intended to provide further explanation without limiting the scope of the disclosure as claimed.
Various embodiments of the present disclosure are discussed in detail below. While specific embodiments are discussed, this is done for illustration purposes only. A person skilled in the relevant art will recognize that other components and configurations may be used without departing from the present disclosure.
As used herein, the terms “first,” “second,” “third,” and “fourth” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.
The terms “forward” and “aft” refer to relative positions within a turbofan engine or vehicle, and refer to the normal operational attitude of the turbofan engine or vehicle. For example, with regard to a turbofan engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.
As used herein, the terms “low,” “mid” (or “mid-level”), and “high,” or their respective comparative degrees (e.g., “lower” and “higher”, where applicable), when used with compressor, combustor, turbine, shaft, fan, or turbofan engine components, each refers to relative pressures, relative speeds, relative temperatures, or relative power outputs within an engine unless otherwise specified. For example, a “low-power” setting defines the engine or the combustor configured to operate at a power output lower than a “high-power” setting of the engine or the combustor, and a “mid-level power” setting defines the engine or the combustor configured to operate at a power output higher than a “low-power” setting and lower than a “high-power” setting. The terms “low,” “mid” (or “mid-level”) or “high” in such aforementioned terms may additionally, or alternatively, be understood as relative to minimum allowable speeds, pressures, or temperatures, or minimum or maximum allowable speeds, pressures, or temperatures relative to normal, desired, steady state, etc., operation of the engine. A mission cycle for a turbofan engine includes, for example, a low-power operation, a mid-level power operation, and a high-power operation. Low-power operation includes, for example, engine start, idle, taxiing, and approach. Mid-level power operation includes, for example, cruise. High-power operation includes, for example, takeoff and climb.
The various power levels of the turbofan engine are defined as a percentage of a sea level static (SLS) maximum engine rated thrust. Low power operation includes, for example, less than thirty percent (30%) of the SLS maximum engine rated thrust of the turbofan engine. Mid-level power operation includes, for example, thirty percent (30%) to eighty-five percent (85%) of the SLS maximum engine rated thrust of the turbofan engine. High power operation includes, for example, greater than eighty-five percent (85%) of the SLS maximum engine rated thrust of the turbofan engine. The values of the thrust for each of the low power operation, the mid-level power operation, and the high power operation of the turbofan engine are exemplary only, and other values of the thrust can be used to define the low power operation, the mid-level power operation, and the high power operation.
The terms “coupled,” “fixed,” “attached,” “connected,” and the like, refer to both direct coupling, fixing, attaching, or connecting, as well as indirect coupling, fixing, attaching, or connecting through one or more intermediate components or features, unless otherwise specified herein.
The singular forms “a,” “an,” and “the” include plural references unless the context clearly dictates otherwise.
As used herein, the terms “axial” and “axially” refer to directions and orientations that extend substantially parallel to a centerline of the turbofan engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the centerline of the turbofan engine. In addition, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations that extend arcuately about the centerline of the turbofan engine.
As used herein, a “turbo-engine” includes a compressor section, a combustion section, and a turbine section.
As used herein, a “turbofan engine” includes a turbo-engine and a fan that directs air into the turbo-engine, and rated for use in a regional aircraft, narrow body aircraft, or wide body aircraft. A turbofan engine rated for use on a regional aircraft will have a maximum takeoff thrust in a range of ten thousand pound-force to twenty thousand pound-force (10,000 lbf to 20,000 lbf). A turbofan engine rated for use on a narrow body aircraft will have a maximum takeoff thrust in a range of fifteen thousand pound-force to thirty thousand pound-force (15,000 lbf to 30,000 lbf). A turbofan engine rated for use on a wide body aircraft will have a maximum takeoff thrust in a range of forty thousand pound-force to one hundred ten thousand pound-force (40,000 lbf to 110,000 lbf).
As used herein, the term “cruise” or “cruising speed” refers to operation of a turbine engine utilized to power an aircraft that may operate at a cruising speed when the aircraft levels after climbing to a specified altitude. A turbine engine may operate at a cruising speed that is from 50% to 90% of a rated speed, such as from 70% to 80% of the rated speed. In some embodiments, a cruising speed may be achieved at about 80% of full throttle, such as from about 50% to about 90% of full throttle, such as from about 70% to about 80% full throttle. As used herein, the term “cruise flight” refers to a phase of flight in which an aircraft levels in altitude after a climb phase and prior to descending to an approach phase. In various examples, cruise flight may take place at a cruise altitude up to approximately 65,000 ft. In certain examples, cruise altitude is between approximately 28,000 ft. and approximately 45,000 ft. In yet other examples, cruise altitude is expressed in flight levels (FL) based on a standard air pressure at sea level, in which cruise flight is between FL280 and FL650. In another example, cruise flight is between FL280 and FL450. In still certain examples, cruise altitude is defined based at least on a barometric pressure, in which cruise altitude is between approximately 4.85 psia and approximately 0.82 psia based on a sea-level pressure of approximately 14.70 psia and sea-level temperature at approximately 59 degrees Fahrenheit. In another example, cruise altitude is between approximately 4.85 psia and approximately 2.14 psia. In certain examples, the ranges of cruise altitude defined by pressure may be adjusted based on a different reference sea-level pressure and/or sea-level temperature.
As used herein, the term “ducted engine” means a turbofan engine with a fan casing or nacelle that circumferentially surrounds the fan.
As used herein, an “unducted fan engine” or an “open fan engine” means a turbofan engine without a fan casing or a nacelle surrounding the fan.
Hereafter, the term “turbofan engine” will refer to either a “ducted engine” or an “open fan engine.”
As used herein, a “fan tip diameter” is defined as a diameter of a fan blade and is measured from the longitudinal centerline axis of the turbofan engine to a fan tip of the fan blade at an axial location of the blade where the diameter is a maximum.
As used herein, a Mach number is a ratio of the speed of the aircraft to the speed of sound in the surrounding airflow. The Mach number at cruise as defined herein is a maximum operating Mach number as provided by a Type Certificate Data Sheet (TCDS) for the turbine engine.
An aircraft's quoted cruise Mach number is generally known in the industry to be applied during a “standard day” temperature day. Therefore, the temperature is a fixed value based on altitude according to the established International Standard Atmosphere (ISA) tables. High speed civil gas turbine powered transport aircraft quote their speed by Mach number and have set cruising altitudes based on their size and mission profile (e.g., smaller aircraft fly at lower altitudes). Turboprops and smaller aircraft may have their cruising speed quoted in knots such as VTAS (velocity true airspeed) or KCAS (knots calibrated air speed), where ambient temperature is considered. Engine performance can be modeled for “hot days” or “cold days” where the ambient temperature is hotter or cooler than standard day by a prescribed amount, but this is part of off-design performance. Further, between 36,000 and 80,000 feet, where most commercial aircraft cruise, the ambient temperature is actually constant.
As used herein, a “thrust bearing radius” of a radial thrust bearing is defined in the radial direction from the longitudinal centerline axis to a radial center of the radial thrust bearing. Particularly, the radial center of the radial thrust bearing is a radial center of the rolling elements of the radial thrust bearing.
As used herein, a “fan hub axial length” is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112) from a fan hub tip of the fan hub to a pitch axis P of the fan blades of the fan.
As used herein, a “fan actuation system axial length” is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface of the fan actuation system to the pitch axis P of the fan blades of the fan.
As used herein, a “fan bearing axial length” is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112) from the pitch axis P of the fan blades of the fan to an axial center of one or more fan bearings that support rotation of the fan shaft.
Here and throughout the specification and claims, range limitations are combined, and interchanged. Such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.
The present disclosure provides for turbofan engines that have a variable pitch fan. Such engines include a fan actuation system that includes one or more actuators for changing a pitch of fan blades of the variable pitch fan. The fan actuation system typically includes a hydraulic system that supplies hydraulic fluid to one or more chambers to actuate the actuators. The actuators are coupled to the fan blades and actuation of the actuators causes the fan blades to rotate about a pitch axis P to change the pitch of the fan blades. Some fan actuation systems are designed for turboprop engines that include a propeller, rather than a fan.
Turboprop engines produce less thrust than turbofan engines. Turboprop engines typically provide cruise speeds for an aircraft with a Mach number that is less than 0.7 and have fewer than ten propeller blades, such as fewer than eight propeller blades or fewer than five propeller blades. Turbofan engines include ten or more fan blades that extend from a disk and provide cruise speeds for an aircraft with a Mach number that is 0.7 or greater. To achieve these higher speeds, the fan aerodynamics for the turbofan engines are different than the propeller aerodynamics for turboprop engines, resulting in the turbofan engines having more fan blades for aerodynamic efficiency at higher Mach speeds. Turbofan engines with variable pitch fan blades also benefit from guide vanes, such as outlet guide vanes behind the fan blades, and/or inlet guide vanes forward of the fan, to reduce losses at higher speeds.
The loading environment associated with the variable pitch mechanism for turboprop engines is less than the loading environment presented for a variable pitch turbofan engine. There is a lower disk loading capability requirement on parts (e.g., trunnion, bearings, gearing, actuators, etc.) and associated less actuation force resources needed (e.g., hydraulic fluid) to operate a variable pitch turboprop as compared to a variable pitch turbofan engine. At the same time, the available space, the desirable space, or the volume in that part of the engine for the higher-load-carrying fan blade pitch actuation system and the greater number of blades of a turbofan engine is not correspondingly larger than the space available for the lower-load-carrying fan blade pitch actuation system with fewer fan blades of a turboprop. Turbofan engines having variable pitch fan blades require more compactness for the pitch change system, relative to a turboprop, when considering the larger space requirements assumed if one were to simply scale-up a pitch actuation system for a turboprop for use in a turbofan engine. This can be realized when one considers that a larger, stronger structure is needed to support the more numerous blades and react the higher pitch loads associated with a turbofan engine. One cannot simply scale-up the space available for a pitch change mechanism and associated structure, and also scale up to account for the impact of a significantly increased number of blades when designing a variable pitch turbofan engine. Accommodation of the pitch change mechanism, trunnion, and associated structure for holding and articulating the fan blades within an engine housing therefore presents unique challenges for the turbofan engine in terms of the available space. The existing pitch change mechanisms and structure used to support blades in turboprop engines are not faced with similar challenges and therefore provide limited insight into how to implement a variable pitch mechanism within the more limited space, and more numerous fan blade system of a turbofan engine.
Many actuation systems for turboprop engines include a counterweight system to help pitch the propeller blades (e.g., the weight counteracts inertial loading associated with turning the propeller blade). For turbofan engines, a counterweight system may not be feasible because there is not the space available to accommodate the counterweight system. Thus, an alternative is needed to articulate the blades without exceeding load limits, which implies more compactness given the limited space available. Additionally, it was realized that pitch lock devices to lock the more-numerous fan blades in a feather position for turbofan engines, in case of fan actuation system failure, need to be considered when determining the minimum size needed for the turbofan engine fan actuation system. Additionally, it should be realized the very different types of inlets between a turboprop engine, on the one hand, and turbofan engine on the other hand, impact the amount of available space within the engine housing. Inlets to the turbo-engine (e.g., the compressor section, the combustion section, and the turbine section) of a turboprop engine have a relatively narrow circumferential extent (sometimes called “chin” inlets). As such, there is more space available for a pitch change mechanism. Inlets to turbofan engines, however, have annular inlets, which take up more space within the engine housing than the more limited circumferential extent occupied by a turboprop inlet. Accommodating both a pitch change mechanism and annular inlet poses a unique challenge for a turbofan engine with variable pitch fan blades.
For at least these reasons, the loading on a pitch change mechanism and packaging of this system for a turbofan engine having greater number of blades than a turboprop engine presents challenges. It is not simply a matter of scaling-up the space available and size of component parts used in a turboprop engine fan actuation system. Indeed, it has been found that the problem is both unique to the engine type and complex—not amenable to a ready solution based on pre-existing variable pitch turboprop engine design. The inventors, seeking a need to find a solution to this problem, designed and tested several different turbofan engine architectures in an effort to arrive at a fan actuation system that met both the higher loading and more compact space requirements of a turbofan engine.
Referring now to the drawings,
The turbo-engine 116 includes, in serial flow relationship, a compressor section 121, a combustion section 126, and a turbine section 127. The turbo-engine 116 depicted is substantially enclosed within a core cowl 118 that is substantially tubular and defines a core inlet 120 having an annular shape that is annular about the longitudinal centerline axis 112. As schematically shown in
For the embodiment depicted in
The gearbox assembly 146 is shown schematically in
Referring still to the exemplary embodiment of
During operation of the turbofan engine 110, a volume of air 158 enters the turbofan engine 110 through an inlet 160 of the nacelle 150 or the fan assembly 114. As the volume of air 158 passes across the fan blades 140, a first portion of air, referred to as bypass air 162, is directed or routed into the bypass airflow passage 156, and a second portion of air, referred to as core air 164, is directed or is routed into the upstream section of the core air flow path, or, more specifically, into the core inlet 120 of the LP compressor 122. The ratio between the bypass air 162 and the core air 164 is commonly known as a bypass ratio. The pressure of the core air 164 is then increased by the LP compressor 122 to form compressed air 165, and the compressed air 165 is routed through the HP compressor 124 and into the combustion section 126, where the compressed air 165 is mixed with fuel and burned to generate combustion gases 166.
The combustion gases 166 are routed into the HP turbine 128 and expanded through the HP turbine 128 where a portion of thermal energy and kinetic energy from the combustion gases 166 is extracted via one or more stages of HP turbine stator vanes 168 that are coupled to the core cowl 118 and HP turbine rotor blades 170 that are coupled to the HP shaft 134. This causes the HP shaft 134 to rotate, thereby supporting operation of the HP compressor 124 (e.g., a self-sustaining cycle). In this way, the combustion gases 166 do work in the HP turbine 128 to cause the HP turbine rotor blades 170 (and the HP shaft 134) to rotate at a sufficient rate to maintain the compression ratio of the HP compressor 124 (e.g., self-sustaining cycle). The combustion gases 166 are then routed into the LP turbine 130 and expanded through the LP turbine 130. Here, a second portion of the thermal energy and the kinetic energy is extracted from the combustion gases 166 via one or more stages of LP turbine stator vanes 172 that are coupled to the core cowl 118 and LP turbine blades 174 that are coupled to the LP shaft 136. This causes the LP shaft 136 to rotate, thereby supporting operation of the LP compressor 122 and rotation of the fan 138 via the gearbox assembly 146 (e.g., a self-sustaining cycle). In this way, the combustion gases 166 do work in the LP turbine 130 to cause the LP turbine blades 174 (and the LP shaft 136) to rotate.
The combustion gases 166 are subsequently routed through the core exhaust nozzle 132 of the turbo-engine 116 to provide propulsive thrust at a thrust level of the turbofan engine 110. The thrust level of the turbofan engine 110 includes a cruise thrust level defined by a cruise Mach number Mcruise that is the Mach number of the turbofan engine 110 at cruise conditions, or mid-level power conditions. Simultaneously, the bypass air 162 is directed through the bypass airflow passage 156 before being exhausted from a fan exhaust nozzle 176 of the turbofan engine 110, also providing propulsive thrust. The HP turbine 128, the LP turbine 130, and the core exhaust nozzle 132 at least partially define a hot gas path 178 for routing the combustion gases 166 through the turbo-engine 116.
The turbofan engine 110 depicted in
As shown in
The turbofan engine 210 includes a turbo-engine 220 and a fan assembly 250 positioned upstream thereof. Generally, the turbo-engine 220 includes a compressor section, a combustion section, a turbine section, and an exhaust section. Particularly, as shown in
The combustion gases flow from the combustor 230 downstream to a high-pressure (HP) turbine 232. The HP turbine 232 drives the HP compressor 228 through a first shaft, also referred to as a high-pressure (HP) shaft 236 (also referred to as a “high-speed shaft”). In this regard, the HP turbine 232 is drivingly coupled with the HP compressor 228. Together, the HP compressor 228, the combustor 230, and the HP turbine 232 define the engine core 218. The combustion gases then flow to a power turbine or a low-pressure (LP) turbine 234. The LP turbine 234 drives the LP compressor 226 and components of the fan assembly 250 through a second shaft, also referred to as a low-pressure (LP) shaft 238 (also referred to as a “low-speed shaft”). In this regard, the LP turbine 234 is drivingly coupled with the LP compressor 226 and components of the fan assembly 250. The LP shaft 238 is coaxial with the HP shaft 236 in the embodiment of
The fan assembly 250 includes a fan 252, also referred to as a primary fan. For the embodiment of
The gearbox assembly 255 is shown schematically in
The fan blades 254 can be arranged in equal spacing around the longitudinal centerline axis 212. Each fan blade 254 extends outwardly from a disk (not shown in
The fan assembly 250 further includes a fan guide vane array 260 that includes a plurality of fan guide vanes 262 (only one shown in
The fan cowl 270 annularly encases at least a portion of the core cowl 222 and is generally positioned outward of the core cowl 222 along the radial direction R. Particularly, a downstream section of the fan cowl 270 extends over a forward portion of the core cowl 222 to define a fan flowpath, also referred to as a fan duct 272. Incoming air enters through the fan duct 272 through a fan duct inlet 276 and exits through a fan exhaust nozzle 278 to produce propulsive thrust. The fan duct 272 is an annular duct positioned generally outward of the core duct 242 along the radial direction R. The fan cowl 270 and the core cowl 222 are connected together and supported by a plurality of struts 274 (only one shown in
The turbofan engine 210 also defines or includes an inlet duct 280. The inlet duct 280 extends between an engine inlet 282 and the core inlet 224 and the fan duct inlet 276. The engine inlet 282 is defined generally at the forward end of the fan cowl 270 and is positioned between the fan 252 and the fan guide vane array 260 along the axial direction A. The inlet duct 280 is an annular duct that is positioned inward of the fan cowl 270 along the radial direction R. Air flowing downstream along the inlet duct 280 is split, not necessarily evenly, into the core duct 242 and the fan duct 272 by a splitter 284 of the core cowl 222. The inlet duct 280 is wider than the core duct 242 along the radial direction R. The inlet duct 280 is also wider than the fan duct 272 along the radial direction R.
The fan assembly 250 also includes a mid-fan 286. The mid-fan 286 includes a plurality of mid-fan blades 288 (only one shown in
Accordingly, air flowing through the inlet duct 280 flows across the plurality of mid-fan blades 288 and is accelerated downstream thereof. At least a portion of the air accelerated by the mid-fan blades 288 flows into the fan duct 272 and is ultimately exhausted through the fan exhaust nozzle 278 to produce propulsive thrust. Also, at least a portion of the air accelerated by the plurality of mid-fan blades 288 flows into the core duct 242 and is ultimately exhausted through the core exhaust nozzle 240 to produce propulsive thrust. Generally, the mid-fan 286 is a compression device positioned downstream of the engine inlet 282. The mid-fan 286 is operable to accelerate air into the fan duct 272, also referred to as a secondary bypass passage.
During operation of the turbofan engine 210, an initial airflow or an incoming airflow passes through the fan blades 254 of the fan 252 and splits into a first airflow and a second airflow. The first airflow bypasses the engine inlet 282 and flows generally along the axial direction A outward of the fan cowl 270 along the radial direction R. The first airflow accelerated by the fan blades 254 passes through the fan guide vanes 262 and continues downstream thereafter to produce a primary propulsion stream or a first thrust stream S1. A majority of the net thrust produced by the turbofan engine 210 is produced by the first thrust stream S1. The second airflow enters the inlet duct 280 through the engine inlet 282.
The second airflow flowing downstream through the inlet duct 280 flows through the plurality of mid-fan blades 288 of the mid-fan 286 and is consequently compressed. The second airflow flowing downstream of the mid-fan blades 288 is split by the splitter 284 located at the forward end of the core cowl 222. Particularly, a portion of the second airflow flowing downstream of the mid-fan 286 flows into the core duct 242 through the core inlet 224. The portion of the second airflow that flows into the core duct 242 is progressively compressed by the LP compressor 226 and the HP compressor 228, and is ultimately discharged into the combustion section. The discharged pressurized air stream flows downstream to the combustor 230 where fuel is introduced to generate combustion gases or products.
The combustor 230 defines an annular combustion chamber that is generally coaxial with the longitudinal centerline axis 212. The combustor 230 receives pressurized air from the HP compressor 228 via a pressure compressor discharge outlet. A portion of the pressurized air flows into a mixer. Fuel is injected by a fuel nozzle (omitted for clarity) to mix with the pressurized air thereby forming a fuel-air mixture that is provided to the combustion chamber for combustion. Ignition of the fuel-air mixture is accomplished by one or more igniters (omitted for clarity), and the resulting combustion gases flow along the axial direction A toward, and into, a first stage turbine nozzle 233 of the HP turbine 232. The first stage turbine nozzle 233 is defined by an annular flow channel that includes a plurality of radially extending, circumferentially spaced nozzle vanes 235 that turn the combustion gases so that the combustion gases flow angularly and impinge upon first stage turbine blades of the HP turbine 232. The combustion gases exit the HP turbine 232 and flow through the LP turbine 234, and exit the core duct 242 through the core exhaust nozzle 240 to produce a core air stream, also referred to as a second thrust stream S2. As noted above, the HP turbine 232 drives the HP compressor 228 via the HP shaft 236, and the LP turbine 234 drives the LP compressor 226, the fan 252, and the mid-fan 286 via the LP shaft 238.
The other portion of the second airflow flowing downstream of the mid-fan 286 is split by the splitter 284 into the fan duct 272. The air enters the fan duct 272 through the fan duct inlet 276. The air flows generally along the axial direction A through the fan duct 272 and is ultimately exhausted from the fan duct 272 through the fan exhaust nozzle 278 to produce a third stream, also referred to as a third thrust stream S3.
The third thrust stream S3 is a secondary air stream that increases fluid energy to produce a minority of total propulsion system thrust. In some embodiments, a pressure ratio of the third stream is higher than that of the primary propulsion stream (e.g., a bypass or a propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of the secondary air stream with the primary propulsion stream or a core air stream, e.g., into a common nozzle. In certain embodiments, an operating temperature of the secondary air stream is less than a maximum compressor discharge temperature for the engine. Furthermore, in certain embodiments, aspects of the third stream (e.g., airstream properties, mixing properties, or exhaust properties), and thereby a percent contribution to total thrust, are passively adjusted during engine operation or can be modified purposefully through the use of engine control features (such as fuel flow, electric machine power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluidic features) to adjust or to improve overall system performance across a broad range of potential operating conditions.
The turbofan engine 210 depicted in
Further, for the depicted embodiment of
In some embodiments, the electric machine 290 can be an electric motor operable to drive or to motor the LP shaft 238. In other embodiments, the electric machine 290 can be an electric generator operable to convert mechanical energy into electrical energy. In this way, electrical power generated by the electric machine 290 can be directed to various engine systems or aircraft systems. In some embodiments, the electric machine 290 can be a motor/generator with dual functionality. The electric machine 290 includes a rotor 294 and a stator 296. The rotor 294 is coupled to the LP shaft 238 and rotates with rotation of the LP shaft 238. In this way, the rotor 294 rotates with respect to the stator 296, thereby generating electrical power. Although the electric machine 290 has been described and illustrated in
The disk 306 includes a plurality of disk segments 312 that is rigidly coupled together or integrally molded together in a generally annular shape. One fan blade 304 is coupled to each disk segment 312 at a trunnion mechanism 314 of the fan actuation system 302. The trunnion mechanism 314 facilitates retaining the respective fan blade 304 on the disk 306 during rotation of the disk 306, while still rendering the respective fan blade 304 rotatable relative to the disk 306 about a pitch axis P of the fan blade 304. For example, the trunnion mechanism 314 provides a load path to the disk 306 for the centrifugal load generated by the fan blade 304 during rotation of the fan blade 304 about the longitudinal centerline axis 301. The trunnion mechanism 314 includes a plurality of bearings disposed within the disk segment 312 that allows the fan blade 304 to rotate about the pitch axis P.
The trunnion mechanism 402 includes a plurality of unison rings 408, 410 including a forward unison ring 408 positioned forward of the plurality of trunnions 404 and an aft unison ring 410 positioned aft of the plurality of trunnions 404. The forward unison ring 408 and the aft unison ring 410 couple the plurality of trunnions 404 together. The plurality of trunnion links 406 is coupled to the forward unison ring 408 or the aft unison ring 410 via a plurality of pins 412. The plurality of forward trunnion links 406a is pivotably coupled to the forward unison ring 408 by a plurality of forward pins 412a such that the plurality of trunnions 404 is coupled to the forward unison ring 408. For example, each forward trunnion link 406a extends forward from a respective trunnion 404 to the forward unison ring 408 and a respective forward pin 412a is disposed through the forward trunnion link 406a at the forward unison ring 408 to pivotably couple the forward trunnion link 406a to the forward unison ring 408. Each aft trunnion link 406b extends aftward from the respective trunnion 404 to the aft unison ring 410 and a respective aft pin 412b is disposed through the aft trunnion link 406b at the aft unison ring 410 to pivotably couple the aft trunnion link 406b to the aft unison ring 410. In this way, each of the plurality of trunnions 404 is pivotably coupled to the forward unison ring 408 and to the aft unison ring 410 such that the plurality of trunnions 404 can pivot about the pitch axis P in unison.
The one or more actuators 414 include a hydraulic cylinder 416 and a piston 418 disposed within the hydraulic cylinder 416. The hydraulic cylinder 416 and the piston 418 are movable along the axial direction A. In this way, the one or more actuators 414 are hydraulic linear actuators such that the hydraulic cylinder 416 and the piston 418 move linearly along the axial direction A (e.g., in opposite directions along the longitudinal centerline axis 112). The forward unison ring 408 is coupled to the hydraulic cylinder 416 such that the forward unison ring 408 moves when the hydraulic cylinder 416 moves. The aft unison ring 410 is coupled to the piston 418 such that aft unison ring 410 moves when the piston 418 moves.
In operation, the fan actuation system 400 moves the plurality of fan blades 140 (
A hydraulic system supplies a hydraulic fluid (e.g., oil) to one or more hydraulic chambers of the one or more actuators 414 to move the hydraulic cylinder 416 and the piston 418 to pitch the plurality of fan blades 140. An exemplary hydraulic system and hydraulic chambers are detailed below with respect to
As the hydraulic cylinder 416 moves axially along the axial direction A, the hydraulic cylinder 416 causes the forward unison ring 408 to move, thereby causing the plurality of forward trunnion links 406a to pivot and to pitch the plurality of trunnions 404, and, therefore, pitching the plurality of fan blades 140 about the pitch axis P. At the same time, movement of the piston 418 along the axial direction A causes the aft unison ring 410 to move, thereby, causing the plurality of aft trunnion links 406b to pivot in an opposite direction as the forward trunnion links 406a, and, therefore, pitching the plurality of fan blades 140 about the pitch axis P. In this way, the fan actuation system 400 translates linear motion of the one or more actuators 414 (e.g., along the axial direction A) into rotational motion of the plurality of fan blades 140. Such a configuration enables a compact and lightweight design of the fan actuation system 400. Further, each of the hydraulic cylinder 416 and the piston 418 provides only half of the force needed to actuate the plurality of trunnions 404 and provides a redundant path in the event that one of the hydraulic cylinder 416 or the piston 418 fails.
The fan actuation system 500 includes a trunnion mechanism 502 including a plurality of trunnions 504. Each fan blade 140 is coupled to a respective one of the plurality of trunnions 504. The plurality of trunnions 504 extends through an opening 505 in the disk 142. The plurality of trunnions 504 is rotatable in the opening 505. This enables the plurality of fan blades 140 to rotate about the pitch axis P. As such, the pitch of the plurality of fan blades 140 can be changed relative to the flow of the volume of air 158. In particular, the plurality of fan blades 140 can be rotated (e.g., pitched) to any position between the first end position (e.g., the feather position) and the second end position (e.g., the reverse position). In
The fan actuation system 500 includes a plurality of trunnion links 506 and a unison ring 508. The plurality of trunnion links 506 is pivotably coupled to the plurality of trunnions 504. For example, each trunnion link 506 is coupled to a respective trunnion 504 and to the unison ring 508. In this way, the unison ring 508 couples the plurality of trunnions 504 together. The plurality of trunnion links 506 is coupled to the unison ring 508 via a plurality of pins 512. In this way, the plurality of trunnions 504 is pivotably coupled to the unison ring 508 such that the plurality of trunnions 504, and, thus, the plurality of fan blades 140, can pivot about the pitch axis P in unison, as detailed further below.
The fan actuation system 500 includes one or more actuators 514 that include a hydraulic cylinder 516, a piston 518, and a piston retainer 520. The piston retainer 520 is coupled (e.g., bolted) to the fan shaft 145 such that the piston retainer 520 rotates with the fan shaft 145. Therefore, the piston retainer 520 is coupled (e.g., indirectly) to, and rotated by, the LP shaft 136 (
In the illustrated example of
The hydraulic cylinder 516 is disposed radially outward of (e.g., around, surrounding) the piston retainer 520 and the piston 518. The hydraulic cylinder 516 is keyed to the piston retainer 520. As such, the piston retainer 520 rotates the hydraulic cylinder 516. However, the hydraulic cylinder 516 is slidable along the piston retainer 520 in the axial direction A (left and right in
The hydraulic cylinder 516 has a first portion 516a, a second portion 516b, a third portion 516c, and a fourth portion 516d. The first portion 516a extends generally in the axial direction A and is coupled to the unison ring 508 at the joint 517 (e.g., a bolted joint). The second portion 516b is disposed radially inward of the first portion 516a and is coupled to the first portion 516a and to the unison ring 508 at the joint 517. The third portion 516c extends forward from the joint 517 (e.g., from the first portion 516a, the second portion 516b, and the unison ring 508) and forms a pressurized pneumatic chamber 570, disclosed in further detail herein. The fourth portion 516d is coupled to, and extends axially within, the third portion 516c. The first portion 516a, the second portion 516b, the third portion 516c, and the fourth portion 516d form the hydraulic cylinder 516. In some examples, the first portion 516a, the second portion 516b, the third portion 516c, and the fourth portion 516d are separate parts or components that are coupled (e.g., welded, bolted) together. In other examples, one or more of the first portion 516a, the second portion 516b, the third portion 516c, and the fourth portion 516d can be constructed as a single unitary part or component (e.g., a monolithic structure). In some embodiments, the hydraulic cylinder 516 and the unison ring 508 form a single unitary part or component.
The first portion 516a of the hydraulic cylinder 516 is sealingly engaged with (e.g., engaged with a seal to prevent fluid leakage) the third portion 520c of the piston retainer 520. The second portion 520b of the piston retainer 520 is sealingly engaged with the first portion 516a of the hydraulic cylinder 516. The second portion 516b of the hydraulic cylinder 516 is sealingly engaged with the first portion 520a of the piston retainer 520. The piston 518 is sealingly engaged with the second portion 516b and with the fourth portion 516d of the hydraulic cylinder 516.
The fan actuation system 500 includes one or more hydraulic chambers defined between the hydraulic cylinder 516, the piston 518, and the piston retainer 520. These hydraulic chamber(s) are used to control the position of the hydraulic cylinder 516, and, thus, to control the pitch of the plurality of fan blades 140. As shown in
The fan actuation system 500 includes a hydraulic system 550 to provide hydraulic fluid, such as oil, to one or more of the hydraulic chambers 540, 542, 544 to control the movement of the hydraulic cylinder 516. The hydraulic system 550 includes a pump 552 to control the first pressure P1 and the second pressure P2. The pump 552 is activated to move the hydraulic fluid into, or out of, the hydraulic chambers 540, 542, 544 to increase or to decrease the first pressure P1 and the second pressure P2, and, therefore, to cause the hydraulic cylinder 516 to move forward or to move rearward. In the illustrated example, the hydraulic system 550 includes an oil transfer bearing 554. The oil transfer bearing 554 includes a fixed portion 556 (e.g., a shaft) with fluid passageways fluidly coupled to the pump 552. The fixed portion 556 is a static component and does not rotate or move axially. The oil transfer bearing 554 includes a sleeve 558 that is rotatable about the fixed portion 556. The hydraulic system 550 includes a first fluid line 560, a second fluid line 562, and a third fluid line 564 fluidly coupled between the oil transfer bearing 554 and the respective hydraulic chambers 540, 542, and 544. The first fluid line 560 is in fluid communication with the first hydraulic chamber 540, the second fluid line 562 is in fluid communication with the second hydraulic chamber 542, and the third fluid line 564 is in fluid communication with the third hydraulic chamber 544. The first fluid line 560, the second fluid line 562, and the third fluid line 564 are coupled to the sleeve 558. The sleeve 558 enables fluid communication among the first fluid line 560, the second fluid line 562, and the third fluid line 564, which are rotating with the fan actuation system 500, and the fixed portion 556 of the oil transfer bearing 554. Thus, the oil transfer bearing 554 enables the hydraulic fluid to be transferred between a stationary component and a rotating component. As disclosed above, the first hydraulic chamber 540 and the third hydraulic chamber 544 are provided with the hydraulic fluid at the same first pressure P1. The oil transfer bearing 554 fluidly couples the hydraulic fluid in the first fluid line 560 and the third fluid lines 564 such that the first hydraulic chamber 540 and the third hydraulic chamber 544 remain at the same first pressure P1. To move the plurality of fan blades 140 away from the feather position and
toward the reverse position, the pump 552 is activated to increase the first pressure P1 in the first hydraulic chamber 540 and the third hydraulic chamber 544 and to reduce the second pressure P2 in the second hydraulic chamber 542. As a result, the hydraulic cylinder 516 moves in the rearward direction (to the right in
The pressurized pneumatic chamber 570 is formed or is defined by the third portion 516c of the hydraulic cylinder 516 and the piston 518. The pressurized pneumatic chamber 570 is filled with a pressurized gas. In some examples, the pressurized pneumatic chamber 570 contains pressurized nitrogen. In other examples, the pressurized pneumatic chamber 570 can be filled with another pressurized gas (e.g., air). The pressurized pneumatic chamber 570 is sealed. A such, the volume of the pressurized gas (e.g., nitrogen) in the pressurized pneumatic chamber 570 does not change. During manufacture or assembly of the fan actuation system 500, the pressurized pneumatic chamber 570 can be charged with gas (e.g., nitrogen) and then sealed. The pressurized pneumatic chamber 570 can be pressurized to any amount depending on the size of the pressurized pneumatic chamber 570 and on the size of the hydraulic chambers 540, 542, 544 and the desired biasing force. In some examples, the pressure in the pressurized pneumatic chamber 570 is in a range of seven hundred twenty pounds per square inch to nine hundred twenty pounds per square inch (720 psi to 920 psi). In other examples, however, the pressure may be less than, or greater than, these exemplary values.
The pressurized gas in the pressurized pneumatic chamber 570 generates a constant force or a constant load that biases the hydraulic cylinder 516 in the forward direction (to the left in
The example pressurized pneumatic chamber 570 is advantageous because it has a high load capability due to the compressibility of the pneumatic gas (e.g., nitrogen). Further, the pressurized pneumatic chamber 570 enables a longer travel of the hydraulic cylinder 516 with relatively little change in load. Therefore, the pressurized pneumatic chamber 570 provides a relatively constant load throughout the stroke. Also, the volume and areas of the pressurized pneumatic chamber 570 and the piston 518 can be varied to optimize the load versus travel of the hydraulic cylinder 516.
Therefore, during normal operation of the fan actuation system 500, the first hydraulic chamber 540 and the third hydraulic chamber 544 act to bias the hydraulic cylinder 516 in the rearward direction, while the second hydraulic chamber 542 and the pressurized pneumatic chamber 570 act to bias the hydraulic cylinder 516 in the forward direction. The pressures in the hydraulic chambers 540, 542, and 544 and in the pressurized pneumatic chamber 570 can be controlled to substantially balance the forces and to maintain the hydraulic cylinder 516 in a desired position. In the illustrated example of
In the example of
Examples have been disclosed herein that improve the ability for the fan actuation system 500 to move the fan blades 140 to the feather position in the event of failure of the fan actuation system 500 or a shutdown of the turbofan engine 110. The example systems disclosed herein are passive and, thus, do not require complicated activation components or control systems. The example pressurized pneumatic chamber 570 is capable of handling high rotational speeds and a large variation in operating temperatures, such as encountered during use on aircraft. The examples disclosed herein also eliminate the need for a pitch lock device. As such, the example systems can result in fewer parts, less complexity, reduced weight, and lower costs compared to known systems. The fan actuation system 500 is particularly useful in turbofan engines (e.g., the turbofan engine 110 of
The turbofan engine 110 also includes one or more thrust bearings, also referred to as one or more radial thrust (radial blade load) bearings 580, disposed between the trunnion 504 and the disk 142 such that the trunnion 504 rotates about the pitch axis P with respect to the disk 142. The one or more radial thrust bearings 580 transmit the load (the radial blade load) from the respective fan blade 140 to a static structure of the turbofan engine 110. In particular, the radial thrust bearings 580 include a plurality of rolling elements 582. The rolling elements 582 can include, for example, ball bearings, tapered roller bearings, or the like, for transmitting the radial blade load from the fan blade 140 to the static structure.
The one or more radial thrust bearings 580 are disposed radially at a thrust bearing radius RTB. The thrust bearing radius RTB is defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 583 of the one or more radial thrust bearings 580. The radial center 583 is a center of the radial thrust bearings 580 in the radial direction R. Particularly, the radial center 583 is defined as a radial center of the rolling elements 582. The amount of space, or the volume, beneath the fan 138 that is available for the fan actuation system 500 is defined by the thrust bearing radius RTB. The fan actuation system 500 needs to be accommodated radially below the one or more radial thrust bearings 580 and within the thrust bearing radius RTB.
The turbofan engine 110 includes a fan hub axial length AFH, a fan actuation system axial length AFAS, and a fan bearing axial length AFB. The fan hub axial length AFH is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112) from the fan hub tip 157 to the pitch axis P of the fan blades 140. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 515 of the fan actuation system 500 to the pitch axis P of the fan blades 140. In
The fan actuation system 600 includes a trunnion mechanism 602, a plurality of trunnions 604, a plurality of trunnion links 606, a unison ring 608, a plurality of pins 612, one or more actuators 614, a hydraulic cylinder 616, a joint 617, a piston 618, and a piston retainer 620. The hydraulic cylinder 616 has a first portion 616a and a second portion 616b. Although not shown in the view of
In the epicyclic gear assembly, the gear assembly 147 can be in a star arrangement or a rotating ring gear type gear assembly (e.g., the third gear 149c is rotating and the planet carrier 151 is fixed and stationary). In such an arrangement, the fan 138 is driven by the third gear 149c. For example, the third gear 149c is coupled to the fan shaft 145 such that rotation of the third gear 149c causes the fan shaft 145, and, thus, the fan 138, to rotate. In this way, the third gear 149c is an output of the gear assembly 147. However, other suitable types of gear assemblies may be employed. In one non-limiting embodiment, the gear assembly 147 is a planetary arrangement, in which the third gear 149c is held fixed, with the planet carrier 151 allowed to rotate. In such an arrangement, the fan 138 is driven by the planet carrier 151. For example, the planet carrier 151 is coupled to the fan shaft 145 such that rotation of the planet carrier 151 causes the fan shaft 145, and, thus, the fan 138, to rotate. In this way, the one or more second gears 149b (e.g., via the planet carrier 151) are the output of the gear assembly 147. In another non-limiting embodiment, the gear assembly 147 may be a differential gear assembly in which the third gear 149c and the planet carrier 151 are both allowed to rotate. While an epicyclic gear assembly is detailed herein, the gear assembly can include any type of gear assembly including, for example, a single stage gear assembly or a compound gear assembly (e.g., a gear assembly having a plurality of stages).
The plurality of gears 149 includes one or more gear bearings 153 disposed therein. For example, the one or more second gears 149b each includes one or more gear bearings 153 disposed therein. The one or more gear bearings 153 enable the plurality of gears 149 to rotate about the one or more gear bearings 153 such that the plurality of gears 149 rotates. The one or more gear bearings 153 can include any type of bearing for a gear, such as, for example, journal bearings, roller bearings, or the like. The gearbox assembly 146 can include a plurality of gear bearings that includes a forward gear bearing and an aft gear bearing. The one or more gear bearings 153 shown in the view of
The first gear 149a is coupled to an input shaft of the turbofan engine 110. For example, the first gear 149a is coupled to the LP shaft 136 such that rotation of the LP shaft 136 causes the first gear 149a to rotate. Radially outward of the first gear 149a, and intermeshing therewith, is the one or more second gears 149b that are coupled together and supported by the planet carrier 151. The planet carrier 151 supports and constrains the one or more second gears 149b such that the each of the one or more second gears 149b is enabled to rotate about a corresponding axis of each second gear 149b without rotating about the periphery of the first gear 149a. Radially outwardly of the one or more second gears 149b, and intermeshing therewith, is the third gear 149c, which is an annular ring gear. The third gear 149c is coupled via an output shaft to the fan 138 and rotates to drive rotation of the fan 138 about the longitudinal centerline axis 112. For example, the fan shaft 145 is coupled to the third gear 149c.
The fan shaft 145 is coupled to the disk 142 such that rotation of the fan shaft 145 causes the plurality of fan blades 140 to rotate about the longitudinal centerline axis 112. The turbofan engine 110 also includes one or more radial thrust bearings 680, disposed between the trunnion 604 and the disk 142 such that the trunnion 604 rotates about the pitch axis P with respect to the disk 142. In particular, the radial thrust bearings 680 include a plurality of rolling elements 682.
The one or more radial thrust bearings 680 are disposed radially at the thrust bearing radius RTB defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 683 of the one or more radial thrust bearings 680, as discussed above. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 615 (shown schematically in
The fan actuation system 700 includes a trunnion mechanism 702, a plurality of trunnions 704, an opening 705, one or more trunnion links 706, a unison ring 708, one or more actuators 714, an axially forward-most surface 715, a piston 718, a piston retainer 720, and one or more radial thrust bearings 780. The piston retainer 720 is stationary (e.g., coupled to a static structure of the turbofan engine 110) and the piston 718 moves with respect to the piston retainer 720 to change a pitch of the fan blades 140. For example, the piston 718 can be coupled to a hydraulic cylinder that receives hydraulic fluid for moving the piston 718, as detailed above. The one or more trunnion links 706 include one or more ring gears that mesh with a corresponding gear of the trunnions 704.
The fan actuation system 700 also includes a counterweight assembly 790 including one or more counterweights 792. The counterweights 792 are axially spaced from the trunnions 704 to counter a centrifugal twisting moment of the fan blades 140. The counterweights 792 can be any high-density mass that can rotate about a counterweight centerline. The counterweights 792 can have offset masses that are movable relative to the counterweight centerline. In particular, the counterweights 792 are coupled to one or more counterweight shafts 794 that are drivingly coupled to the trunnion links 706 via one or more counterweight gears 795. The counterweight shafts 794 are supported by one or more counterweight support members 796 that are coupled to the piston retainer 720. In
As the trunnions 704 rotate, the trunnions 704 cause the trunnion links 706 to rotate with respect to the unison ring 708, and in turn, the trunnion links 706 cause the counterweight shafts 794 to rotate. As the trunnion links 706 and the counterweight shafts 794 rotate, the counterweights 792 rotate via the counterweight shafts 794. In this way, the counterweights 792 change position relative to the counterweight centerline. Thus, the counterweight assembly 790 counters a centrifugal twisting moment of the fan blades 140 to help rotate the fan blades 140 when the pitch of the fan blades 140 changes.
A mass of the counterweights 792 can be changed based on a length of the counterweight shafts 794. In particular, the counterweights 792 can have less mass with longer counterweight shafts 794 and can have more mass with shorter counterweight shafts 794. In this way, the axially further the counterweights 792 are disposed from the pitch axis P of the fan blades 140, the lesser mass the counterweights 792 can have, while still countering the centrifugal twisting moment of the fan blades 140 and helping to rotate the fan blades 140 when the pitch of the fan blades 140 changes. Accordingly, the mass of the counterweights 792 needed to pitch the fan blades 140 and counter the twisting moment is a function of the axial position of the counterweights 792 with respect to the pitch axis P.
The one or more radial thrust bearings 780 are disposed radially at the thrust bearing radius RTB defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 783 of a plurality of rolling elements 782 of the radial thrust bearings 780, as discussed above. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 715 of the fan actuation system 700 to the pitch axis P of the fan blades 140.
The fan actuation system 800 includes a trunnion mechanism 802, a plurality of trunnions 804, an opening 805, one or more trunnion links 806, a plurality of pins 812, one or more actuators 814 (shown schematically in
The counterweight assembly 890 includes one or more counterweights 892, one or more counterweight shafts 894, and one or more counterweight support members 896. The one or more counterweight support members 896 are coupled to the disk 142 such that the counterweight assembly 890 rotates about the longitudinal centerline axis 112 with rotation of the fan 138. The counterweight assembly 890 also includes one or more link arms 895 and one or more lever arms 898. The one or more lever arms 898 are pivotably coupled to the counterweight support members 896 via a pivot 899. The link arms 895 are coupled to the trunnion links 806 via the pins 812 and are pivotably coupled to the lever arms 898. The counterweight shafts 894 are pivotably coupled to the lever arms 898 at the pivot 899.
In
As the trunnions 804 rotate, the trunnions 804 cause the trunnion links 806 to rotate, and in turn, the trunnion links 806 cause the pins 812 to rotate, and, thus, cause the link arms 895 to pivot. As the link arms 895 pivot, the link arms 895 cause the lever arms 898 to pivot, and, thus, cause the counterweight shafts 894 to pivot about the pivot 899. In this way, the counterweight shafts 894 cause the counterweights 892 to travel along a partially circular arc radially outward away from the longitudinal centerline axis 112 or radially inward towards the longitudinal centerline axis 112. Thus, the counterweight assembly 890 counters a centrifugal twisting moment of the fan blades 140 to help rotate the fan blades 140 when the pitch of the fan blades 140 changes.
The one or more radial thrust bearings 880 are disposed radially at the thrust bearing radius RTB defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 883 of a plurality of rolling elements 882 of the radial thrust bearings 880, as discussed above. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 815 of the fan actuation system 800 to the pitch axis P of the fan blades 140.
The fan actuation system 900 includes a trunnion mechanism 902, a plurality of trunnions 904, an opening 905, one or more trunnion links 906, a unison ring 908, one or more actuators 914, an axially forward-most surface 915, and one or more radial thrust bearings 980. The actuators 914 can include any of the actuators disclosed herein for changing a pitch of the fan blades 140. The one or more trunnion links 906 and the unison ring 908 couple the trunnions 904 to the actuators 914 such that movement of the actuators 914 causes the trunnions 904 to rotate, thus, causing the fan blades 140 to rotate about the pitch axis P.
The counterweight assembly 990 includes one or more counterweights 992, one or more counterweight shafts 994, one or more counterweight support members 996, and one or more lever arms 998. In
The counterweight assembly 990 includes a counterweight hub 997 that may be connected to the disk 142, such that rotation of the disk 142 about the longitudinal centerline axis 112 drives rotation of the counterweight hub 997 about the longitudinal centerline axis 112. The counterweight shafts 994 are rotationally connected to the counterweight hub 997. For example, each of the counterweight shafts 994 may be mounted to the counterweight hub 997 via one or more bearing members 993 that provide the ability for the counterweight shafts 994 to rotate about a counterweight lever rotational axis PCW. The counterweight bearings 993 may be any type of bearing (e.g., tapered roller bearings, spherical roller bearings, cylindrical roller bearings, needle roller bearings, thrust ball bearings, angular contact roller bearings, deep groove ball bearings, etc.), and are not limited to any particular type of bearing Each of the counterweight support members 996 are rotational about a counterweight lever rotational axis PCW that extends through a respective counterweight support member 996 and extends radially (i.e., in the radial direction R) from the longitudinal centerline axis 112.
Each counterweight shaft 994 is a cantilever arm having a first end connected to a respective counterweight support member 996 and a second end offset from the respective counterweight lever rotational axis PCW. A respective counterweight 992 is connected to the second end of the counterweight shaft 994. Each counterweight 992 has a counterweight center-of-gravity that is utilized in locating the counterweight 992 within the counterweight assembly 990.
The one or more counterweight support members 996 are coupled to the disk 142 such that the counterweight assembly 990 rotates about the longitudinal centerline axis 112 with rotation of the fan 138. The counterweight assembly 990 also includes one or more lever arms 998 that are rotationally connected to the actuators 914 via one or more lever bearings 999. The lever arms 998 are connected to the counterweight support members 996 such that axial translation of the actuators 914 along the longitudinal centerline axis 112 drives the lever arms 998 and the counterweight support members 996 about the respective counterweight lever rotational axis PCW so as to rotate the counterweight shafts 994. In
In
As the actuators 914 move axially, the actuators 914 cause the trunnions 904 and the counterweight support members 996 to rotate. In turn, the counterweight support members 996 cause the counterweight shafts 994 to rotate about the counterweight lever rotational axis PCW, and, thus, cause the counterweights 992 to rotate. In particular, the counterweight shafts 994, and the counterweights 992, rotate in to or out of the page between the ninety-degree rotated position that defines a maximum axial extent of the counterweights 992 and a zero-degree rotated position that defines a minimum axial extend of the counterweights 992. Thus, the counterweight assembly 990 counters a centrifugal twisting moment of the fan blades 140 to help rotate the fan blades 140 when the pitch of the fan blades 140 changes.
The one or more radial thrust bearings 980 are disposed radially at the thrust bearing radius RTB defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 983 of a plurality of rolling elements 982 of the radial thrust bearings 980, as discussed above. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 915 of the fan actuation system 900 to the pitch axis P of the fan blades 140.
The fan actuation system 1000 includes a trunnion mechanism 1002, a plurality of trunnions 1004, an opening 1005, one or more trunnion links 1006, a unison ring 1008, one or more actuators 1014, an axially forward-most surface 1015, one or more radial thrust bearings 1080, and a counterweight assembly 1090. The actuators 1014 can include any of the actuators disclosed herein for changing a pitch of the fan blades 140. The one or more trunnion links 1006 and the unison ring 1008 couple the trunnions 1004 to the actuators 1014 such that movement of the actuators 1014 causes the trunnions 1004 to rotate, thus, causing the fan blades 140 to rotate about the pitch axis P. In
The counterweight assembly 1090 includes one or more counterweights 1092, one or more counterweight shafts 1094, and one or more counterweight support members 1096. The one or more counterweight support members 1096 are coupled to the disk 142 via the unison ring 1008 such that the counterweight assembly 1090 rotates about the longitudinal centerline axis 112 with rotation of the fan 138. The counterweights 1092 are positioned axially aft of the fan blades 140, particularly, axially aft of the pitch axis P. For example, the counterweights 1092 are positioned axially between the pitch axis P and the fan bearings 155.
The counterweight support members 1096 act as a carrier for the counterweight shafts 1094. The counterweight shafts 1094 are aligned generally parallel to the longitudinal centerline axis 112 and pass through the counterweight support members 1096. The counterweight shafts 1094 are rotatably connected (e.g., via one or more gears) at a first end to the unison ring 1008. The counterweights 1092 are connected to a second end of the counterweight shafts 1094. The counterweight shafts 1094, and the counterweights 1092, are rotatable relative to the counterweight support members 1096, about a respective counterweight shaft axis PCWS.
All of the counterweight shafts 1094 are meshed via one or more gears with the unison ring 1008. Thus connected, the movement of the fan blades 140, unison ring 1008, and the counterweights 1092 are linked together such that rotary motion of the unison ring 1008, for example, caused by the actuators 1014, will cause a simultaneous change in the pitch angle of all of the fan blades 140, and of the angular orientation of the counterweights 1092. The unison ring 1008 transmits forces between the fan blades 140 and the counterweights 1092. In this way, the counterweight shafts 1094 cause the counterweights 1092 to travel along a partially circular arc radially outward away from the longitudinal centerline axis 112 or radially inward towards the longitudinal centerline axis 112, and axially closer to, or axially further from, the pitch axis P. Thus, the counterweight assembly 1090 counters a centrifugal twisting moment of the fan blades 140 to help rotate the fan blades 140 when the pitch of the fan blades 140 changes.
The one or more radial thrust bearings 1080 are disposed radially at the thrust bearing radius RTB defined in the radial direction R from the longitudinal centerline axis 112 to a radial center 1083 of a plurality of rolling elements 1082 of the radial thrust bearings 1080, as discussed above. The fan actuation system axial length AFAS is an axial length, in the axial direction (e.g., parallel with the longitudinal centerline axis 112), from an axially forward-most surface 1015 of the fan actuation system 1000 to the pitch axis P of the fan blades 140.
As mentioned earlier, the inventors sought to address the problem implementing a variable pitch actuation system within the more limited packaging space available in a turbofan engine and while accounting for the significantly higher loading environment and more numerous blades relative to a turboprop engine. By way of testing various engine architectures the inventors experimented with different configurations of the pitch actuation system, fine and coarse pitch actuators, hydraulic actuators, and bearing placement that could sustain the higher loading associated with more numerous blades, higher disk loading, and Mach speed sufficient to satisfy operational and safety requirements in the event of, e.g., loss of hydraulic pressure. Additionally, while it was possible to arrive at such a system after experiments and testing, there was a challenge to determine how to fit the system within a comparatively more limited space of a turbofan engine.
During the course of evaluating the different embodiments as set forth herein, with the goal of providing the necessary force to pitch the fan blades, taking due account for the number of blades, accounting for loss in fluid pressure or generally lost power conditions, aerodynamic performance, cooling, aeromechanics, and disc loading/fan blade loading, etc., the inventors had discovered there was indeed much less space available for this system to operate as required for the engine's pitch actuation system. After evaluating several different architectures of pitch change mechanisms (with and without counterweight, oil transfer devices, fine and coarse pitch system, torque transfer load path for pitching blades and delivery of shaft power from gearbox, etc.—both for a ducted engine and an open fan engine—it was discovered, unexpectedly, that there is relationships among the number of fan blades, the fan tip diameter DFT, the cruise Mach number, and the thrust bearing radius RTB, and an axial length LAXIAL capable of differentiating an architecture that satisfies operational and packaging requirements from an architecture that does not satisfy these requirements. These relationships moreover are capable of uniquely identifying a finite and readily ascertainable number of embodiments suitable for a particular architecture that accounts for the size and the loading requirements needed to pitch the fan blades without overly sacrificing the aerodynamic performance, cooling aeromechanics, and load margins on the fan blades. For example, the cruise Mach number was not expected to be a significant factor, but as discussed further below, the cruise Mach number was found to be a factor and particularly in conjunction with fan diameter at higher Mach numbers. The inventors submit that the relationships enable one to select a size for the fan pitch actuation system that can reduce the size and the weight of the fan pitch actuation system, while accounting for the factors discussed above. The inventors further submit that the relationships can help identify an improved fan efficiency, or penalties to efficiency by choosing one fan pitch actuation system architecture over another. A relationship is referred to as a fan actuation system (FAS) envelope, in relationship (1):
NFB is the number of fan blades of the fan, DFT is the fan tip diameter, Mcruise is the Mach number at cruise (mid-level power operation), and RTB is the thrust bearing radius of the radial thrust bearings (any of the radial thrust bearings detailed herein). NFB×DFT×Mcruise is referred to as a loading envelope, and RTB/NFB is referred to as a spacing envelope. Accordingly, the FAS envelope is given by the loading envelope divided by the spacing envelope.
A second relationship is referred to as a fan actuation system length (FASL) envelope, in relationship (2):
NFB is the number of fan blades of the fan, DFT is the fan tip diameter, RTB is the thrust bearing radius of the radial thrust bearings, and LAXIAL is an axial length, along the longitudinal centerline axis 112 from the fan hub tip 157 to the fan bearings 155. In particular, LAXIAL is a summation of the fan hub axial length AFH and the fan bearing axial length AFB. NFB×DFT is referred to as a loading envelope, and LAXIAL×(RTB/NFB) is referred to as a spacing envelope. Accordingly, the FASL envelope is given by the loading envelope divided by the spacing envelope.
As discussed further below, the inventors identified a range for the FAS envelope and the FASL envelope that enables a fan actuation system design for different turbofan engine architectures that accounts for the integrity/reliability of load paths needed to pitch the fan blades within the space constraints imposed by a turbofan engine (vs. a turboprop's space constraints). Fan pitch actuation system architectures that fall within this range are believed to satisfy packaging requirements for a turbofan engine, while those architectures that do not fall within the FAS envelope range or the FASL envelope range are believed to not satisfy the packaging requirements, which indicate that the system would be unacceptably large and not result in an aircraft engine that met aero efficiency and weight requirements (i.e., an undesirable engine architecture). Using these unique relationships, the size of the fan actuation system can be selected to achieve a more compact fan pitch actuation system for a turbofan engine. Using the FAS envelope or the FASL envelope as a guide, a fan pitch actuation system can be developed that takes into account the loading associated with pitching of the fan blades based on the size of the fan blades, the number of fan blades, the size of thrust bearing, the cruise Mach number, or the axial length, which factors were found—as a result of the extensive number of architectures considered for different thrust class engines, some successful and some not successful—to largely define the packaging size needed to accommodate a pitch actuation system capable of handling the fan loading environment.
Table 1 represents exemplary embodiments 1 to 14 and their corresponding FAS envelope and FASL envelope values for various turbofan engines at various cruise Mach numbers. Embodiments 1 to 14 may represent the turbofan engine 110 of
The FAS envelope and the FASL envelope are only valid for an engine with fan blades NFB in a range of from ten to eighteen for a ducted engine, and from ten to fourteen for an open fan engine. The number of fan blades NFB affects the volume (e.g., amount of space) circumscribed by the fan blades. Increasing the number of fan blades NFB increases the amount of airflow that the fan can produce for a particular fan tip diameter and fan rotation speed, but a higher NFB also reduces the tangential distance TFB between fan blades at the fan hub, which impacts the available space for pitch actuation of each individual blade, referring to the space needed per blade for pitch levers, gearing, oil transfer devices, related mechanisms for pitching fan blades and size of load bearing parts of the trunnion and related supporting structure capable of carrying the fan blade loads. This space is at a premium because with an increased number of fan blades the loading capability per blade needs to be satisfied within a smaller space compared to an engine with fewer blades (e.g., such as a turboprop engine). The FAS envelope values and the FASL envelope values account for the number of fan blades NFB selected to increase the amount of airflow but without imposing an unrealistically narrow tangential fan blade distance TEB between adjacent fan blades in order to fit within the desired packaging envelope.
The FAS envelope and the FASL envelope are only valid for a fan tip diameter DFT in a range of eighty-four inches to one hundred eighty inches (84.0 in. to 180.0 in.). In some embodiments, the FAS envelope and the FASL envelope are valid for a fan tip diameter DFT in a range of eighty-four inches to one hundred sixty-eight inches (84.0 in. to 168.0 in.). The fan tip diameter DFT also affects the volume needed for supporting the fan blades during operation. Increasing the fan tip diameter DFT increases the fan tip speed for a given rotational speed and therefore the load that needs to get reacted at the trunnion, and torque needed in the pitching mechanism for pitching the blade. The radial spacing between blades and within the volume circumscribed by the fan blades (e.g., within the space circumscribed by the radial thrust bearings) decreases, thereby decreasing the volume beneath the fan and providing less space for the load bearing structure that can react the blade loads. Furthermore, as the bearing radius RTB is extended out, the structure supporting the blade at its root needs to be capable of sustaining higher loads because the blade is disposed further from the fan rotation axis. The more robust root results in a larger fan disk, further providing less space underneath the fan for the fan actuation system. In view of these weight and size considerations, as well as the ability to install such fan blades and fans without resulting in unacceptable aero efficiency penalties, the inventors determined that a fan tip diameter DFT should be less than one hundred eighty inches (180.0 in.). In some embodiments, the fan tip diameter DFT should be less than one hundred sixty-eight inches (168.0 in.). The fan tip diameter DFT may therefore be limited as it impacts the space available for a pitch actuation system suitable for carrying fan blade loads. The size of the fan blades in ducted engines is limited by the duct (e.g., the nacelle). In embodiments for a ducted engine (e.g., the turbofan engine 110 of
The FAS envelope and the FASL envelope are only valid for a thrust bearing radius RTB in a range of twelve inches to twenty-seven inches (12 in. to 27 in.). In some embodiments, the thrust bearing radius RTB is in a range of fourteen inches to twenty-seven inches (14 in. to 27 in.). The thrust bearing radius RTB defines the amount of space, or the volume available for the fan actuation system. Increasing the thrust bearing radius RTB provides more space for the fan actuation system, but sacrifices aerodynamic performance by making the fan radius ratio (i.e., the ratio of the fan hub radius to the fan blade radius) larger. Decreasing the thrust bearing radius RTB reduces the fan radius ratio and reduces the size of the turbofan engine, but provides less space to carry the loads from the fan blades. The thrust bearing radius RTB reflects the need for adequately accommodating the diameter needed for packaging the fan actuation system but without overly sacrificing aerodynamic performance of the turbofan engine. In embodiments for a ducted engine (e.g., the turbofan engine 110 of
The FAS envelope and the FASL envelope are valid for a cruise Mach number Mcruise in a range of 0.7 to 0.92. In some embodiments, the FAS envelope and the FASL envelope are valid for a cruise Mach number Mcruise in a range of 0.7 to 0.9. As mentioned above, turbofan engines operate at higher cruise speeds than turboprop engines. At higher cruise speeds, the aerodynamic loads on fan blades increase, thereby requiring more torque for actuating blades in pitch. This means a larger actuation system is needed to handle the higher reaction loads resulting when a torque is applied in flight to change the blade pitch, to move the blade to a feathered position, or coarse/fine pitch changes. The cruise Mach number Mcruise reflects this higher loading environment when pitching fan blades. In some embodiments, the cruise Mach number Mcruise in a range of 0.75 to 0.9. In some embodiments, the cruise Mach number Mcruise is in a range of 0.8 to 0.88.
The FAS envelope and the FASL envelope are only valid for a fan hub axial length AFH of twenty-five inches to seventy-five inches (25 in. to 75 in.). The fan hub axial length AFH defines the amount of axial space, or the volume available for the fan actuation system, forward of the pitch axis P of the fan blades 140. Increasing the fan hub axial length AFH provides more space for the fan actuation system but increases the overall weight of the turbine engine. Decreasing the fan hub axial length AFH reduces the fan performance and the pressure distribution to the fan due to a smaller axial length for the aerodynamic flow lines into the fan hub but provides less axial space to fit the fan actuation system within the fan hub 148. The fan hub axial length AFH reflects the need for aerodynamic performance for the fan and adequately accommodating the axial length needed for packaging the fan actuation system but without overly sacrificing aerodynamic performance of the turbofan engine and allowing for a more efficient fan actuation system. In embodiments for a ducted engine (e.g., the turbofan engine 110 of
The FAS envelope and the FASL envelope are only valid for a fan bearing axial length AFB of sixteen inches to twenty-three inches (16 in. to 23 in.). The fan bearing axial length AFB defines the amount of axial space, or the volume available for the fan actuation system, aft of the pitch axis P of the fan blades 140. Increasing the fan bearing axial length AFB provides more space for the fan actuation system but increases the overall weight of the engine and increases loads on the bearings. Decreasing the fan bearing axial length AFB decreases overall engine weight and reduces loads on the bearings but provides less axial space to fit the fan actuation system within the fan hub 148. The fan bearing axial length AFB reflects the need for adequately accommodating the axial length needed for packaging the fan actuation system while minimizing the fan bearing axial length AFB to reduce loads on the bearings and reduce overall weight of the engine. In embodiments for a ducted engine (e.g., the turbofan engine 110 of
A first area 1102 represents the boundaries of the FAS envelope for ducted engines, such as, for example, the turbofan engine 110 of
A turbofan engine for an aircraft, the turbofan engine comprising a fan having a plurality of fan blades, each of the plurality of fan blades being rotatable about a pitch axis, and a fan actuation system including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system envelope in a range of 300 to 1860, the fan actuation system envelope being given by:
A turbofan engine for an aircraft, the turbofan engine comprising a fan having a plurality of fan blades, each of the plurality of fan blades being rotatable about a pitch axis, a nacelle that circumferentially surrounds the fan, and a fan actuation system including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system envelope in a range of 300 to 660, the fan actuation system envelope being given by:
A turbofan engine for an aircraft, the turbofan engine comprising a fan having a plurality of fan blades, the fan being an open fan, and each of the plurality of fan blades being rotatable about a pitch axis, and a fan actuation system including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system envelope in a range of 660 to 1860, the fan actuation system envelope being given by:
The turbofan engine of any preceding clause, wherein the fan actuation system includes a hydraulic system that supplies hydraulic fluid for rotating the plurality of fan blades about the pitch axis.
The turbofan engine of any preceding clause, wherein the cruise operating conditions occur at a mid-level power range of the turbine engine.
The turbofan engine of the preceding clause, wherein the mid-level power range is 30% to 85% of a sea level static maximum engine rated thrust for the turbine engine.
The turbofan engine of any preceding clause, wherein the turbofan engine is rated for use on a regional aircraft having a maximum takeoff thrust of 10,000 lbf to 20,000 lbf.
The turbofan engine of any preceding clause, wherein the turbofan engine is rated for use on a narrow body aircraft having a maximum takeoff thrust of 15,000 lbf to 30,000 lbf.
The turbofan engine of any preceding clause, wherein the turbofan engine is rated for use on a wide body aircraft having a maximum takeoff thrust of 40,000 lbf to 110,000 lbf.
The turbofan engine of any preceding clause, wherein NFB is in a range of ten to eighteen.
The turbofan engine of any preceding clause, wherein NFB is in a range of ten to fourteen.
The turbofan engine of any preceding clause, wherein DFT is in a range of 84.0 inches to 180.0 inches.
The turbofan engine of any preceding clause, wherein DFT is in a range of 84.0 inches to 120.0 inches.
The turbofan engine of any preceding clause, wherein DFT is in a range of 120.0 inches to 168.0 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 14 inches to 27 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 14 inches to 19 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 19 inches to 27 inches.
The turbofan engine of any preceding clause, wherein Mcruise is in a range of 0.7 to 0.92.
The turbofan engine of any preceding clause, wherein Mcruise is in a range of 0.75 to 0.9.
The turbofan engine of any preceding clause, wherein Mcruise is in a range of 0.8 to 0.88.
The turbofan engine of any preceding clause, wherein the fan actuation system includes a pressurized pneumatic chamber that is filled with a pressurized gas that biases the plurality of fan blades to a feather position.
The turbofan engine of any preceding clause, wherein the fan actuation system is devoid of counterweights for reducing inertial loading associated with rotation of fan blades.
The turbofan engine of any preceding clause, further comprising a turbo-engine, wherein the turbofan engine has a longitudinal centerline axis, and the turbo-engine is annular about the longitudinal centerline axis.
The turbofan engine of the preceding clause, wherein the turbo-engine includes a core inlet that is annular about the longitudinal centerline axis.
The turbofan engine of any preceding clause, further comprising a gearbox assembly, wherein the turbo-engine includes a low-pressure shaft, and the fan has a fan shaft that is coupled to the low-pressure shaft through the gearbox assembly.
The turbofan engine of the preceding clause, wherein the gearbox assembly has a gear ratio in a range 3.5:1 to 5:1 for a ducted engine.
The turbofan engine of any preceding clause, wherein the gearbox assembly has a gear ratio in a range of 4:1 and 10:1 for an unducted fan engine.
The turbofan engine of any preceding clause, wherein the low-pressure shaft, the gearbox assembly, and the fan shaft are coaxial along the longitudinal centerline axis.
The turbofan engine of any preceding clause, wherein the fan actuation system envelope is in a range of 660 to 1020.
The turbofan engine of any preceding clause, the fan actuation system envelope being in a range of 300 to 660.
The turbofan engine of any preceding clause, the fan actuation system envelope being in a range of 660 to 1860.
The turbofan engine of any preceding clause, the fan actuation system envelope being in a range of 660 to 1020.
The turbofan engine of any preceding clause, further comprising a nacelle that circumferentially surrounds the fan.
The turbofan engine of any preceding clause, wherein the turbofan engine is an open fan engine.
The turbofan engine of any preceding clause, further comprising a fan hub, the plurality of fan blades extending radially from the fan hub.
The turbofan engine of any preceding clause, the fan actuation system being disposed within the fan hub.
The turbofan engine of any preceding clause, the turbo-engine comprising a compressor section, a combustor, and a turbine section.
The turbofan engine of any preceding clause, the compressor section including a low-pressure compressor and a high-pressure compressor, and the turbine section including a high-pressure turbine and a low-pressure turbine.
The turbofan engine of any preceding clause, further comprising a high-pressure shaft that couples the high-pressure compressor and the high-pressure turbine.
The turbofan engine of any preceding clause, further comprising a low-pressure shaft that couples the low-pressure compressor and the low-pressure turbine.
The turbofan engine of any preceding clause, the low-pressure shaft being disposed through the high-pressure shaft.
The turbofan engine of any preceding clause, the gearbox assembly comprising a gear assembly comprising a plurality of gears.
The turbofan engine of any preceding clause, the gearbox assembly including one or more gear bearings.
The turbofan engine of any preceding clause, each of the plurality of fan blades extending from a fan root to a fan tip.
The turbofan engine of any preceding clause, the fan tip diameter DFT being defined from the longitudinal centerline axis to the fan tip of each of the plurality of fan blades.
The turbofan engine of any preceding clause, the fan actuation system including a trunnion mechanism that includes a plurality of trunnions, each fan blade being disposed in a respective trunnion.
The turbofan engine of any preceding clause, the fan blades extending from a disk.
The turbofan engine of any preceding clause, the disk including a plurality of disk segments.
The turbofan engine of any preceding clause, each fan blade being coupled to a respective disk segment at the trunnion mechanism.
The turbofan engine of any preceding clause, the plurality of trunnions being rotatable to rotate the plurality of fan blades about the pitch axis.
The turbofan engine of any preceding clause, the fan actuation system including one or more actuators coupled to the plurality of trunnions.
The turbofan engine of any preceding clause, the fan actuation system including a plurality of trunnion links and a unison ring, the plurality of trunnion links being coupled to the plurality of trunnions and to the unison ring.
The turbofan engine of any preceding clause, the plurality of trunnion links including a plurality of forward trunnion links and a plurality of aft trunnion links.
The turbofan engine of any preceding clause, the unison ring including a plurality of unison rings including a forward unison ring that is positioned forward of the plurality of trunnions and an aft unison ring that is disposed aft of the plurality of trunnions.
The turbofan engine of any preceding clause, the plurality of forward trunnion links being coupled to the forward unison ring.
The turbofan engine of any preceding clause, the plurality of aft trunnion links being coupled to the aft unison ring.
The turbofan engine of any preceding clause, further comprising a plurality of pins that couple the plurality of trunnion links to the unison ring.
The turbofan engine of any preceding clause, the plurality of forward trunnion links being coupled to the forward unison ring by a plurality of forward pins.
The turbofan engine of any preceding clause, the plurality of aft trunnion links being coupled to the aft unison ring by a plurality of aft pins.
The turbofan engine of any preceding clause, the one or more actuators including a hydraulic cylinder and a piston disposed within the hydraulic cylinder.
The turbofan engine of the preceding clause, the hydraulic cylinder and the piston being movable along an axial direction.
The turbofan engine of any preceding clause, the forward unison ring being coupled to the hydraulic cylinder such that the forward unison ring moves when the hydraulic cylinder moves.
The turbofan engine of any preceding clause, the aft unison ring being coupled to the piston such that the aft unison ring moves as the piston moves.
The turbofan engine of any preceding clause, the fan actuation system rotating the plurality of fan blades between a first end position and a second end position.
The turbofan engine of any preceding clause, the first end position being a feather position in which the plurality of fan blades is substantially aligned with a flow of a volume of air across the plurality of fan blades.
The turbofan engine of the preceding clause, the fan actuation system rotating the plurality of fan blades to any position between the first end position and the second end position.
The turbofan engine of any preceding clause, the second end positioned being a reverse position in which the plurality of fan blades exceeds a plane that is transverse to the longitudinal centerline axis by at least 30° to assist with braking the aircraft.
The turbofan engine of any preceding clause, the fan actuation system moving the hydraulic cylinder in a first direction and moving the piston in a second direction.
The turbofan engine of any preceding clause, movement of the hydraulic cylinder and the piston causing the plurality of fan blades to rotate about the pitch axis.
The turbofan engine of any preceding clause, the one or more actuators including a piston retainer.
The turbofan engine of the preceding clause, the piston retainer being coupled to the fan shaft such that the piston retainer rotates with the fan shaft.
The turbofan engine of any preceding clause, the piston being coupled to the piston retainer such that the piston rotates with the piston retainer.
The turbofan engine of any preceding clause, the hydraulic cylinder being axially slidable with respect to the piston and the piston retainer.
The turbofan engine of any preceding clause, the piston retainer comprising a first portion, a second portion that extends radially outward from the first portion, and a third portion that extends axially from the second portion.
The turbofan engine of any preceding clause, the third portion of the piston retainer being coupled to the fan shaft.
The turbofan engine of any preceding clause, the piston being coupled to, and extending forward from, the first portion of the piston retainer.
The turbofan engine of any preceding clause, the hydraulic cylinder being disposed radially outward of the piston retainer and the piston.
The turbofan engine of any preceding clause, the hydraulic cylinder being coupled to the unison ring at a joint such that movement of the hydraulic cylinder in the axial direction causes the plurality of fan blades to pitch about the pitch axis.
The turbofan engine of any preceding clause, the hydraulic cylinder having a first portion, a second portion, a third portion, and a fourth portion.
The turbofan engine of the preceding clause, the first portion of the hydraulic cylinder extending generally in the axial direction and being coupled to the unison ring at the joint.
The turbofan engine of any preceding clause, the second portion of the hydraulic cylinder being disposed radially inward of the first portion and being coupled to the first portion and to the unison ring at the joint.
The turbofan engine of any preceding clause, the third portion of the hydraulic cylinder extending forward from the joint.
The turbofan engine of any preceding clause, the fourth portion of the hydraulic cylinder being coupled to, and extending axially within, the third portion of the hydraulic cylinder.
The turbofan engine of any preceding clause, the first portion of the hydraulic cylinder being sealingly engaged with the third portion of the piston retainer.
The turbofan engine of any preceding clause, the second portion of the piston retainer being sealingly engaged with the first portion of the hydraulic cylinder.
The turbofan engine of any preceding clause, the piston being sealingly engaged with the second portion and the fourth portion of the hydraulic cylinder.
The turbofan engine of any preceding clause, the fan actuation system including one or more hydraulic chambers defined between the hydraulic cylinder, the piston, and the piston retainer.
The turbofan engine of the preceding clause, the one or more hydraulic chambers including a first hydraulic chamber, a second hydraulic chamber, and a third hydraulic chamber.
The turbofan engine of any preceding clause, the first hydraulic chamber being defined between first portion of the hydraulic cylinder, the second portion of the piston retainer, and the third portion of the piston retainer.
The turbofan engine of any preceding clause, the second hydraulic chamber being defined between the first portion of the hydraulic cylinder, the second portion of the hydraulic cylinder, the first portion of the piston retainer, and the second portion of the piston retainer.
The turbofan engine of any preceding clause, the third hydraulic chamber being defined between the second portion of the hydraulic cylinder, an aft end of the piston, and the first portion of the piston retainer,
The turbofan engine of any preceding clause, the first hydraulic chamber and the third hydraulic chamber being supplied with a hydraulic fluid at a first pressure, and the second hydraulic chamber being supplied with the hydraulic fluid at a second pressure.
The turbofan engine of any preceding clause, the first pressure and the second pressure being increased or decreased to cause the hydraulic cylinder to move axially forward or axially rearward to rotate the plurality of fan blades about the pitch axis.
The turbofan engine of any preceding clause, the fan actuation system comprising a hydraulic system that supplies the hydraulic fluid to the one or more hydraulic chambers.
The turbofan engine of any preceding clause, the hydraulic system including a pump to supply the hydraulic fluid to the one or more hydraulic chambers.
The turbofan engine of the preceding clause, the hydraulic system comprising an oil transfer bearing including a fixed portion with a plurality of fluid lines coupled to the pump.
The turbofan engine of the preceding clause, the oil transfer bearing including a sleeve that is rotatable about the fixed portion.
The turbofan engine of any preceding clause, the plurality of fluid lines including a first fluid line in fluid communication with the first hydraulic chamber, a second fluid line in fluid communication with the second hydraulic chamber, and a third fluid line in fluid communication the third hydraulic chamber.
The turbofan engine of any preceding clause, the plurality of fluid lines being coupled to the sleeve.
The turbofan engine of any preceding clause, the first hydraulic chamber and the third hydraulic chamber being provided with the hydraulic fluid at the same first pressure.
The turbofan engine of any preceding clause, the pump supplying the hydraulic fluid to the first hydraulic chamber and the third hydraulic chamber to increase the first pressure P1 and supplying the hydraulic fluid to the second hydraulic chamber to decrease the second pressure P2, to move the hydraulic cylinder in the rearward direction to rotate the plurality of fan blades towards the reverse position.
The turbofan engine of any preceding clause, the pump supplying the hydraulic fluid to the second hydraulic chamber to increase the second pressure P2 and supplying the hydraulic fluid to the first hydraulic chamber and the third hydraulic chamber to decrease the first pressure P1, to move the hydraulic cylinder in the forward direction to rotate the plurality of fan blades towards the feather position.
The turbofan engine of any preceding clause, the one or more actuators further comprising a pressurized pneumatic chamber filled with a pressurized gas to bias the hydraulic cylinder to move the plurality of fan blades to the feather position.
The turbofan engine of any preceding clause, a pressure of the pressurized gas in the pressurized pneumatic chamber being in a range of 720 psi to 920 psi.
The turbofan engine of any preceding clause, the pressurized gas in the pressurized pneumatic chamber causing the hydraulic cylinder to move rearward when the hydraulic system or the turbofan engine fails or is shut down.
The turbofan engine of any preceding clause, the fan actuation system not including a pitch lock device.
The turbofan engine of any preceding clause, the one or more radial thrust bearings being disposed between the plurality of trunnions and the disk such that the plurality of trunnions rotates with respect to the disk to rotate the plurality of fan blades about the pitch axis.
The turbofan engine of any preceding clause, the one or more radial thrust bearings transmitting a load from the plurality of fan blades to a static structure of the turbofan engine.
A turbofan engine for an aircraft, the turbofan engine including a fan having a plurality of fan blades coupled to a fan shaft having one or more fan bearings, each of the plurality of fan blades being rotatable about a pitch axis and extending from a fan hub, and a fan actuation system disposed within the fan hub and including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24, the fan actuation system length envelope being given by
wherein NFB is a number of the plurality of fan blades, DFT is a fan tip diameter of the plurality of fan blades, RTB is a thrust bearing radius of the one or more radial thrust bearings, and LAXIAL is an axial length from a fan hub tip of the fan hub to the one or more fan bearings.
The turbofan engine of the preceding clause, wherein the fan actuation system includes a pressurized pneumatic chamber that is filled with a pressurized gas that biases the plurality of fan blades to a feather position.
The turbofan engine of any preceding clause, wherein the fan actuation system includes one or more counterweights for reducing inertial loading associated with rotation of the plurality of fan blades.
The turbofan engine of any preceding clause, further comprising a turbo-engine, wherein the turbofan engine has a longitudinal centerline axis, and the turbo-engine is annular about the longitudinal centerline axis wherein the turbo-engine includes a core inlet that is annular about the longitudinal centerline axis.
The turbofan engine of any preceding clause, wherein the fan actuation system includes a hydraulic system that supplies hydraulic fluid for rotating the plurality of fan blades about the pitch axis.
The turbofan engine of any preceding clause, wherein NFB is in a range of ten to eighteen.
The turbofan engine of any preceding clause, wherein NFB is in a range of ten to fourteen.
The turbofan engine of any preceding clause, wherein DFT is in a range of 84.0 inches to 180.0 inches.
The turbofan engine of any preceding clause, wherein DFT is in a range of 84.0 inches to 120.0 inches.
The turbofan engine of any preceding clause, wherein DFT is in a range of 120.0 inches to 180.0 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 12 inches to 27 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 12 inches to 19 inches.
The turbofan engine of any preceding clause, wherein RTB is in a range of 19 inches to 27 inches.
The turbofan engine of any preceding clause, wherein LAXIAL is given by AFH+AFB, AFH being a fan hub axial length from the fan hub tip to the pitch axis of the plurality of fan blades and AFB being a fan bearing axial length from the pitch axis of the plurality of fan blades to the one or more fan bearings.
The turbofan engine of any preceding clause, wherein AFH is in a range of 25 inches to 75 inches.
The turbofan engine of any preceding clause, wherein AFB is in a range of 16 inches to 23 inches.
The turbofan engine of any preceding clause, wherein the fan actuation system has a fan actuation system axial length (AFAS) defined from an axially forward-most surface of the fan actuation system to the pitch axis of the plurality of fan blades, AFAS being a maximum of 80% AFH.
A turbofan engine for an aircraft, the turbofan engine including a fan having a plurality of fan blades coupled to a fan shaft having one or more fan bearings, each of the plurality of fan blades being rotatable about a pitch axis and extending from a fan hub, a nacelle that circumferentially surrounds the fan, and a fan actuation system disposed within the fan hub and including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 13, the fan actuation system length envelope being given by
wherein NFB is a number of the plurality of fan blades, DFT is a fan tip diameter of the plurality of fan blades, RTB is a thrust bearing radius of the one or more radial thrust bearings, and LAXIAL is an axial length from a fan hub tip of the fan hub to the one or more fan bearings, wherein LAXIAL is given by AFH+AFB, AFH being a fan hub axial length from the fan hub tip to the pitch axis of the plurality of fan blades and AFB being a fan bearing axial length from the pitch axis of the plurality of fan blades to the one or more fan bearings, AFH is in a range of 25 inches to 40 inches, and AFB is in a range of 17 inches to 20 inches.
A turbofan engine for an aircraft, the turbofan engine including a fan having a plurality of fan blades coupled to a fan shaft having one or more fan bearings, the fan being an open fan, and each of the plurality of fan blades being rotatable about a pitch axis and extending from a fan hub, and a fan actuation system disposed within the fan hub and including one or more actuators for rotating the plurality of fan blades about the pitch axis and one or more radial thrust bearings, wherein the fan actuation system is characterized by a fan actuation system length envelope in a range of 8.5 to 24, the fan actuation system length envelope being given by
wherein NFB is a number of the plurality of fan blades, DFT is a fan tip diameter of the plurality of fan blades, RTB is a thrust bearing radius of the one or more radial thrust bearings, and LAXIAL is an axial length from a fan hub tip of the fan hub to the one or more fan bearings, and RTB is a thrust bearing radius of the one or more radial thrust bearings, wherein LAXIAL is given by AFH+AFB, AFH being a fan hub axial length from the fan hub tip to the pitch axis of the plurality of fan blades and AFB being a fan bearing axial length from the pitch axis of the plurality of fan blades to the one or more fan bearings, AFH is in a range of 25 inches to 75 inches, and AFB is in a range of 16 inches to 23 inches, and DFT is in a range of 120.0 inches to 180.0 inches.
The turbofan engine of the preceding clause, wherein RTB is in a range of 12 inches to 27 inches.
Although the foregoing description is directed to the preferred embodiments of the present disclosure, other variations and modifications will be apparent to those skilled in the art and may be made without departing from the disclosure. Moreover, features described in connection with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.
This application is a continuation-in-part of U.S. patent application Ser. No. 18/400,746, filed on Dec. 29, 2023, which is hereby incorporated by reference herein in its entirety.
| Number | Date | Country | |
|---|---|---|---|
| Parent | 18400746 | Dec 2023 | US |
| Child | 19097493 | US |
