System and method for reducing lash in engine

Abstract

A method for reducing a clearance between a plurality of rotor blades and a shroud assembly of an engine includes determining, with a flight control system, that an aircraft is in a first flight condition. The method also includes adjusting the gap to a first gap distance associated with the first flight condition. Further, the method includes receiving, with the flight control system, a demand for a second flight condition. The method includes adjusting at least two independently controllable parameters during a second flight condition, the at least two independently controllable parameters including at least a first parameter for optimizing the gap and a second parameter for satisfying a thrust demand of the engine, the first parameter having a first influence on the gap and the second parameter having a second influence on the gap, the first impact is greater than the second impact.

Description

Technical Field

[0001] The present disclosure relates generally to aircraft engines and, more particularly, to controlling clearances within an aircraft engine so that the aircraft engine operates more efficiently during operation. Background Art

[0002] At least some known aircraft include an engine control system, sometimes referred to as a full authority digital engine control (FADEC). A FADEC is a system that includes a digital computer and its associated accessories for controlling various aspects of aircraft engine performance. The FADEC receives multiple current input variables for current flight conditions, including but not limited to current values ​​of air density, throttle lever position, engine temperature, engine pressure, and other engine parameters. The inputs are received and analyzed multiple times per second. From this data, engine operating parameters, such as fuel flow, stator vane position, exhaust valve position, etc., are calculated and applied as appropriate to provide optimal engine efficiency given the current flight conditions.

[0003] An aircraft also typically includes a flight control system, which may include a system commonly referred to as a flight management system (FMS). The FMS is a dedicated computer system that automates various flight tasks, including in-flight management of the flight plan. Using various sensors, such as, but not limited to, the Global Positioning System (GPS), Inertial Navigation System (INS), and supplemented by radio navigation to determine the aircraft's position, the FMS guides the aircraft along the flight plan. From the cockpit, the FMS is typically controlled through a control display unit (CDU), which contains a small screen and a keyboard or touch screen. The FMS transmits the flight plan for display on the EFIS, navigation display (ND), or multi-function display (MFD). The FADEC and FMS are independent systems that may transmit current values ​​of parameters in some cases.

[0004] Some known aircraft engines include a hot section module and a cold section module, which include a compressor and a turbine. In order to improve fuel efficiency, thrust and / or turbine life, at least some known engines attempt to control the distance or gap between the tip of each turbine or compressor blade and the surrounding stator. However, due to increases in both rotational speed and operating temperature, the blade tip length measured from the center of the rotor may increase. Both of these effects may be caused by increased fuel flow during maneuvers such as certain actions in a climb, descent / landing sequence, and / or evasive actions. In addition, during operation, especially during transient operations, such as those requiring increased fuel flow, the blade tip length may expand faster than the shroud or stator. Therefore, during such operation, the blade tip may contact the shroud, a condition known as rubbing.

[0005] At least some known aircraft engines use active clearance control to prevent rubbing while minimizing operating clearances under all operating conditions. In at least some conventional systems, active clearance control attempts to expand or contract the shroud or stator by changing the thermal environment of the hardware. However, these systems rely on hardware time constants, and transient lags are typically much slower than the mechanical growth of the hardware due to speed changes.

[0006] Additionally, at least some known aircraft engines activate clearance control in response to one or more engine operating parameters. Additionally, at least some known aircraft engines activate clearance control based on elapsed time relative to transient engine conditions (e.g., throttle manipulation and / or rotor speed changes). Additionally, at least some known aircraft engines adjust clearance control based on, for example, aircraft altitude. Other known active clearance controls are based on mathematical models that are based on data acquired from one or more aircraft engines.

[0007] However, such control may not adequately anticipate increases in fuel flow to mitigate blade tip rubbing due to insufficient clearance. For example, during flight when a throttle change is required to climb from one altitude to another, aircraft engine response is often increased based on a predetermined schedule, which causes the rotor blades to grow (e.g., lengthen) faster than the shrouds surrounding them due to mechanical acceleration of the rotor blades. As engine speed increases, the clearance control system lags behind the relatively rapid expansion of the blades. Therefore, there is a need in the art for improved control strategies to prevent rubbing. BRIEF DESCRIPTION OF THE DRAWINGS

[0008] A complete and enabling disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0009] Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure.

[0010] Figure 2 yes Figure 1 An enlarged cross-sectional schematic illustration of a portion of a gas turbine engine including a clearance control system is shown.

[0011] Figure 3 yes Figure 2 An enlarged cross-sectional schematic illustration of a portion of a gas turbine engine is shown.

[0012] Figure 4 is a schematic block diagram of a system for reducing a turbine or compressor clearance between a plurality of rotor blades of a turbine engine and a shroud of the turbine engine according to an exemplary embodiment of the present invention.

[0013] Figure 5is a flow chart of a method for reducing a turbine clearance between a plurality of rotor blades of a turbine engine and a shroud of the turbine engine according to an exemplary embodiment of the invention.

[0014] Fig. 6A is a graph showing a comparison of a current thrust response of an aircraft engine and the thrust response of an aircraft engine using a clearance control system according to an embodiment of the present disclosure.

[0015] Figure 6B is a graph showing a comparison of a current fan speed response of an aircraft engine and the fan speed of an aircraft engine using a clearance control system according to an embodiment of the present disclosure.

[0016] Figure 6C is a graph showing a comparison of turbine clearance response of an aircraft engine and turbine clearance of an aircraft engine using a clearance control system according to an embodiment of the present disclosure.

[0017] Fig.6D is a graph showing a comparison of the current fan pitch response of an aircraft engine and the fan pitch of an engine using a clearance control system according to an embodiment of the present disclosure.

[0018] Figure 7 is a flow chart of an embodiment of a method for controlling a clearance between a plurality of rotor blades and a shroud of an aircraft engine according to an embodiment of the present disclosure.

[0019] Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present disclosure. DETAILED DESCRIPTION

[0020] Reference will now be made in detail to embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. Each example is provided in a manner to explain the present disclosure, rather than to limit the present disclosure. In fact, it will be appreciated by those skilled in the art that various modifications and variations may be made to the present disclosure without departing from the scope or spirit of the present disclosure. For example, a feature illustrated or described as part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, the present disclosure is intended to encompass such modifications and variations within the scope of the appended claims and their equivalents.

[0021] As used herein, the word "exemplary" means "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or better than other implementations. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0022] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0023] For example, the term “at least one” in the context of “at least one of A, B, and C” means only A, only B, only C, or any combination of A, B, and C.

[0024] The term “turbomachine” refers to a machine that includes one or more compressors, a heat generating section (eg, a combustion section), and one or more turbines that together produce a torque output.

[0025] The term "gas turbine engine" refers to an engine having a turbine as all or part of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojets, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.

[0026] The term "combustion section" refers to any heat addition system of a turbomachine. For example, the term "combustion section" may refer to a section that includes one or more of a deflagration combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or combinations thereof.

[0027] As used herein, the term "rotor" refers to any component of a rotating machine (eg, a turbine engine) that rotates about an axis of rotation. For example, the rotor may include a shaft or spool of a rotating machine (eg, a turbine engine).

[0028] As used herein, the term "stator" refers to any component of a rotating machine (e.g., a turbine engine) having a configuration and arrangement coaxial with a rotor of the rotating machine. The stator may be arranged radially inward or radially outward along a radial axis relative to at least a portion of the rotor. Additionally or alternatively, the stator may be axially arranged adjacent to at least a portion of the rotor.

[0029] The terms "low" and "high" or their respective comparatives (e.g., lower, higher, where applicable), when used with a compressor, turbine, shaft or spool component, etc., refer to relative speeds within an engine unless otherwise specified. For example, a "low turbine" or "low-speed turbine" defines a component that is configured to operate at a lower speed (e.g., maximum allowable speed) than a "high turbine" or "high-speed turbine" of the engine.

[0030] The terms "front" and "rear" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, for a gas turbine engine, front refers to a position closer to the engine inlet, while rear refers to a position closer to the engine nozzle or exhaust.

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

[0032] As used herein, the terms "axial" and "axially" refer to directions and orientations extending substantially parallel to the centerline of the gas turbine engine. Additionally, the terms "radial" and "radially" refer to directions and orientations extending substantially perpendicular to the centerline of the gas turbine engine. Additionally, as used herein, the terms "circumferential" and "circumferentially" refer to directions and orientations extending in an arc around the centerline of the gas turbine engine.

[0033] The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0034] As used herein, the terms "first," "second," "third," etc. may be used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0035] The term "adjacent" as used herein with respect to two walls and / or surfaces means that the two walls and / or surfaces are in contact with each other, or that the two walls and / or surfaces are separated only by one or more non-structural layers, and the two walls and / or surfaces are in a series contact relationship with the one or more non-structural layers (i.e., a first wall / surface contacts the one or more non-structural layers, and the one or more non-structural layers contacts the second wall / surface).

[0036] As used herein, the term "integral" or "unitary" used to describe a structure refers to a structure formed integrally from a continuous material or group of materials without seams, connecting joints, etc. The integral, unitary structures described herein can be formed by additive manufacturing to have the structure, or formed by a casting process, etc.

[0037] Approximate language used in the entire specification and claims is used to modify any quantitative representation that can allow variation without causing the basic function associated therewith to change. Therefore, the value modified by one or more terms (such as "about", "approximately" and "substantially") is not limited to the specified exact value. In at least some cases, approximate language can correspond to the accuracy of the instrument used to measure the value, or the accuracy of the method or machine used to construct or manufacture parts and / or systems. For example, approximate language can refer to within a margin of 1%, 2%, 4%, 10%, 15% or 20%. These approximate margins can be applicable to a single value, any one or two endpoints of a defined numerical range, and / or the margin of the range between the endpoints.

[0038] Here and throughout the specification and claims, range limitations are combinable and interchangeable, and these ranges are identified and include all subranges contained therein unless context or language dictates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints are independently combinable with each other.

[0039] The present disclosure generally relates to systems and methods for reducing the turbine or compressor clearance between a plurality of rotor blades of a turbine engine and a shroud or stator of the turbine engine. For example, in an embodiment, a control logic can be activated (e.g., turned on) in conjunction with adjusting the fan pitch and controlling the fan speed to minimize or otherwise reduce the transient turbine or compressor clearance effect without sacrificing thrust response. This can be achieved by controlling the rate of change of the fan speed to minimize or otherwise reduce the clearance reduction, and by controlling the thrust at a given fan speed using the fan pitch. In addition, in an embodiment, the fan pitch and fan speed can be independently controlled to meet the total fan thrust requirement. In addition, during this transient period, the fan pitch can be maintained below aerodynamic and / or operability limits.

[0040] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, gas turbine engine 10 is a high bypass turbofan jet engine, sometimes also referred to as a "turbofan engine". Figure 1 As shown, the gas turbine engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference), a radial direction R, and a circumferential direction C extending about the longitudinal centerline 12. Generally, the gas turbine engine 10 includes a fan section 14 and a turbine 16 downstream of the fan section 14.

[0041] The illustrated turbine 16 generally includes a tubular outer casing 18 defining an annular inlet 20. The outer casing 18 encloses, in serial flow relationship: a compressor section including a supercharger or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and an injection exhaust nozzle section 32. A high pressure (HP) shaft 34 (which may additionally or alternatively be a spool) drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft 36 (which may additionally or alternatively be a spool) drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and injection exhaust nozzle section 32 together define a working gas flow path 37.

[0042] For the illustrated embodiment, the fan section 14 includes a fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operably coupled to a pitch change mechanism 44, the pitch change mechanism 44 being configured to jointly change the pitch of the fan blades 40, for example, uniformly change the pitch of the fan blades 40. The gas turbine engine 10 may also include a power gearbox 46, with the fan blades 40, the disk 42, and the pitch change mechanism 44 rotating together about the longitudinal centerline 12 via the LP shaft 36 across the power gearbox 46. The power gearbox 46 includes a plurality of gears for adjusting the rotational speed of the fan 38 relative to the rotational speed of the LP shaft 36, so that the fan 38 can rotate at a more efficient fan speed. Thus, the illustrated gas turbine engine 10 is configured as both a geared gas turbine engine (ie, includes a power gearbox 46 ) and a variable pitch gas turbine engine (ie, includes a fan 38 configured as a variable pitch fan).

[0043] Still refer to Figure 1 In the embodiment of the present invention, the disk 42 is covered by a rotatable front hub 48 (sometimes also referred to as a "spinner") of the fan section 14. The front hub 48 has an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and / or the turbine 16. It should be understood that in the illustrated embodiment, the nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the nacelle 50 extends over an outer portion of the turbine 16, thereby defining a bypass airflow passage 56 therebetween.

[0044] During operation of the gas turbine engine 10, a volume of air 58 enters the gas turbine engine 10 through the nacelle 50 and the associated inlet 60 of the fan section 14. As the volume of air 58 passes through the fan blades 40, a first portion of air 62 is directed or channeled into the bypass airflow passage 56, while a second portion of air 64 is directed or channeled into the working gas flow path 37, or more specifically, into the LP compressor 22, as indicated by arrow 64. The ratio between the first portion of air 62 and the second portion of air 64 is generally referred to as a bypass ratio. As the second portion of air 64 is directed through the HP compressor 24 and into the combustion section 26, its pressure increases, where the second portion of air 64 is mixed with fuel and combusted in the combustion section 26 to provide combustion gases 66.

[0045] The combustion gases 66 are directed through the HP turbine 28 where a portion of the thermal and / or kinetic energy in the combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft 34, thereby rotating the HP shaft 34, which supports operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30 where a second portion of the thermal and kinetic energy in the combustion gases 66 is extracted via successive stages of LP turbine stator blades 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thereby rotating the LP shaft 36, which supports operation of the LP compressor 22 and / or rotation of the fan 38.

[0046] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased, also providing propulsive thrust, as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the gas turbine engine 10. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine 16.

[0047] However, it should be understood that Figure 1 The gas turbine engine 10 shown is only an example, and in other exemplary embodiments, the gas turbine engine 10 may have any other suitable configuration. For example, although the gas turbine engine 10 shown is configured as a ducted gas turbine engine (i.e., including an outer nacelle 50), in other embodiments, the gas turbine engine 10 may be a non-ducted gas turbine engine (such that the fan 38 is a non-ducted fan and the outlet guide vanes 52 are cantilevered from the outer casing 18). It should also be understood that in other embodiments, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine. For example, in other embodiments, aspects of the present disclosure may be incorporated into, for example, a turboprop gas turbine engine, a turboshaft gas turbine engine, a turbojet gas turbine engine, or a hybrid electric engine (as the case may be). In additional embodiments, aspects of the present disclosure may be incorporated into an engine that burns hydrogen, sustainable aviation fuel, or is driven by a hydrogen fuel cell.

[0048] Reference now Figure 2 and 3 , shows an enlarged cross-sectional schematic diagram of a portion of a gas turbine engine 10 including a clearance control system 150. In one embodiment, as Figure 2As shown, the HP turbine 28 is substantially coaxially coupled to the HP compressor 24 and the combustion section 26 and is located downstream thereof. The HP turbine 28 includes a rotor assembly 80, which includes at least one rotor 82 formed by one or more disks 84. In an embodiment, the disks 84 include an outer edge 86 and an integral web 88, which extends generally radially therebetween and radially inward from a corresponding blade dovetail slot 90. Each disk 84 also includes a plurality of HP turbine rotor blades 70 extending radially outward from the outer edge 86. The disks 84 include a rear surface 92 and an upstream surface 94.

[0049] like Figure 3 As shown, the annular shroud assembly 96, also referred to as a static casing assembly, surrounds a row of HP turbine rotor blades 70 and maintains a close clearance relationship therewith. In an embodiment, the shroud assembly 96 extends radially inward from the surrounding turbine casing 98 and includes a plurality of shroud members 100 ( Figure 3 ) or arc sectors. In addition, as shown, adjacent shroud components 100 are coupled together so that the shroud components 100 surround the HP turbine rotor blades 70.

[0050] Each shroud member 100 includes a radially outer surface 102 and an opposing radially inner surface 104. A gap 106 is defined between the shroud inner surface 104 and a blade tip 108 of the HP turbine rotor blade 70. More specifically, the gap 106 is defined as the distance between the blade tip 108 and the shroud inner surface 104. It should also be understood that although the present disclosure is described with reference to the HP turbine 28, the gas turbine engine 10 may include multiple stages including multiple rotor blades and gaps associated with each engine module and stage.

[0051] Thus, the clearance control system 150 helps control the clearance 106 during engine operation. Figure 4 , shows a schematic block diagram of an embodiment of a clearance control system 150 according to the present disclosure. In an embodiment, the clearance control system 150 includes a controller 152, such as but not limited to a FADEC. The controller 152 includes a processor 156 and a memory 158 communicatively coupled to the processor 156. The turbine engine 10 described herein is represented as an engine 154 having a fan 160 and a core engine 162 in serial flow communication. In some embodiments, almost all of the airflow through the fan 160 passes through the core engine 162. In various embodiments, the engine 154 is a high bypass engine, and only a portion of the airflow entering the fan 160 passes through the core engine 162. Although described as a FADEC, in various embodiments, the controller 152 may include other forms of engine controllers capable of operating as described herein.

[0052] A plurality of process sensors 164 are positioned about the engine 154 to sense process parameters associated with the engine 154. These process parameters may include, for example, engine speed, fuel flow, damper and guide vane positions, stator vane clearance, and various temperatures of components in the engine 154. The process sensors 164 are communicatively coupled to the controller 152.

[0053] The clearance control system 150 may also be communicatively coupled to a flight control system 170 (e.g., a flight management system or FMS), such as through a communication channel 172. In addition, as shown, the flight control system 170 includes a processor 174 and a memory 176 communicatively coupled to the processor 174. In an embodiment, the communication channel 172 is a wired connection between the controller 152 and the flight control system 170. In other embodiments, the communication channel 172 may be a wireless communication medium. In an embodiment, the flight control system 170 is located near a cockpit (not shown) of an aircraft, and the controller 152 is located near an engine associated therewith. The flight control system 170 may be embodied in a single processor-based component, or the functions of the flight control system 170 may be performed by multiple components configured to perform the functions described herein. Some components that perform the functions of the flight control system 170 may be located near the cockpit, while other components may be distributed within the aircraft for convenience, safety, and / or optimal operation. Although the flight control system is described herein as a flight management system (FMS), it should be understood that the systems and methods described herein include communications between engine controllers and any avionics functions installed on the aircraft.

[0054] The flight control system 170 is configured to interface with various other systems on and off the aircraft. For example, the flight control system 170 may receive current aircraft status from a plurality of aircraft sensors 178 via a sensing system 180. Such aircraft sensors 178 may include a pitot tube for determining airspeed, a gyroscope, a compass, an accelerometer, a position sensor, an altimeter, and various other sensors capable of detecting the condition, status, or position of the aircraft.

[0055] In addition, if Figure 4 As shown, the flight control system 170 may also receive information from one or more onboard processing systems 182, which may be stand-alone systems or systems with functionality distributed across multiple computer systems. In addition, the flight control system 170 and the onboard processing system 182 may use wired communication channels and / or network connections (e.g., Ethernet or fiber optic), wireless communication methods (e.g., radio frequency (RF), such as FM radio and / or digital audio broadcasting, Institute of Electrical and Electronics Engineers, etc.), and / or wireless communication methods (e.g., radio frequency (RF), such as FM radio and / or digital audio broadcasting, Institute of Electrical and Electronics Engineers, etc.). 802.11 standards (e.g., 802.11(g) or 802.11(n)), Worldwide Interoperability for Microwave Access standards, cellular telephone technology (e.g., Global Standard for Mobile Communications (GSM)), satellite communication links, etc.), and / or any other suitable communication means.

[0056] As used herein, wired communication channels include channels that communicate using fiber optics and other optical devices. Flight control system 170 may also receive information from one or more off-board processing systems 184, which may be stand-alone systems or systems with functionality distributed across multiple computer systems and / or multiple sites. Off-board processing system 184 and flight control system 170 are communicatively coupled using one or more wireless communication media, including but not limited to radio frequency (RF), such as FM radio and / or digital audio broadcasting, Institute of Electrical and Electronics Engineers (IEEE) wireless communications. 802.11 standards (e.g. 802.11(g) or 802.11(n)), Worldwide Interoperability for Microwave Access standards, cellular telephone technology (such as the Global Standard for Mobile Communications (GSM)), satellite communication links, and / or any other suitable communication means.

[0057] As with at least some known aircraft operating procedures, a step climb maneuver occurs when the pilot of the aircraft chooses to increase the altitude at which the aircraft is flying. Under current FAA regulations, altitude steps are typically in increments of 2,000 feet. This means, for example, that a pilot of an aircraft flying at an altitude of 33,000 feet may choose to perform a step climb maneuver to cause the aircraft to climb 2,000 feet to an altitude of 35,000 feet. To achieve the step climb maneuver, the pilot modifies the control of the autopilot / autothrottle system of the flight control system 170 to request that the aircraft rise to a desired cruising altitude. The flight control system 170 then uses a predetermined algorithm to increase engine power to cause the aircraft to climb. Since the request for increased engine power generally requires the gas turbine engine 10 to rotate faster to increase engine thrust, the HP turbine rotor blades 70 will grow due to mechanical forces and associated thermal changes. This turbine blade growth can result in, for example, reduced clearance within the HP turbine 28. If this growth exceeds the design clearance, the HP turbine rotor blades 70 will rub against the turbine casing 98 of the gas turbine engine 10, thereby potentially damaging engine components or reducing engine efficiency.

[0058] For example, a step climb of an aircraft from approximately 33,000 feet to 35,000 feet may take more than two minutes to complete. However, known flight control system step climb algorithms command the gas turbine engine 10 to respond to a request for increased thrust within, for example, 5 seconds, resulting in a growth rate of the HP turbine rotor blades 70 that exceeds the growth rate of the engine 10 casing. Since the HP turbine rotor blades 70 grow faster than the surrounding turbine casing 98, engine designers must consider additional clearances to prevent rubbing in these situations. In the above example, the additional clearance is referred to herein as step climb clearance. However, by increasing the clearance between the end of the HP turbine rotor blade 70 and the turbine casing 98, more air is able to escape through the HP turbine rotor blade 70 instead of passing through the HP turbine rotor blade 70, which can result in reduced engine performance and increased fuel consumption. Therefore, it is necessary to develop an algorithm that takes into account controlling (e.g., minimizing) the clearance 106 during step climb transient events.

[0059] The function of the clearance control system 150 is to adjust the active clearance control cooling flow, for example, by increasing the cooling flow to retract the turbine shroud to reduce the clearance in cruise mode, and reducing the cooling flow during step climb to increase the clearance so that no rubbing occurs when the fan speed reaches high values. Due to the thermal time constant of the turbine shroud, the clearance does not increase instantaneously, but gradually so that when the fan speed accelerates to high values, the clearance is large enough to prevent rubbing.

[0060] Reference now Figure 5 , showing a reduction in the engine (eg, engine 10 (eg Figure 1 Flowchart of an embodiment of a method 200 for determining a gap between a plurality of rotor blades and a shroud (shown). Generally, the method 200 will be described herein with reference to the engine 10 described herein. However, it should be understood that the disclosed method 200 may be implemented with any engine having any other suitable configuration. In addition, although Figure 5 The steps performed in a particular order are depicted for the purpose of illustration and discussion, but the methods discussed herein are not limited to any particular order or arrangement. It will be understood by those skilled in the art using the disclosure provided herein that the various steps of the methods disclosed herein may be omitted, rearranged, combined, and / or adapted in various ways without departing from the scope of the present disclosure.

[0061] As shown at (202), method 200 includes determining, using a flight control system (e.g., flight control system 170), that an aircraft is in a first flight condition. For example, in an embodiment, the first flight condition may correspond to a steady-state cruise condition, for example, where cooling flow is increased to contract the shroud assembly 96 to achieve tight clearance to improve engine performance. More specifically, in an embodiment, to determine whether the aircraft is in a cruise state, the engine control system 152 may check the following parameters: the aircraft's cruise altitude is greater than 29,000 feet; the cruise altitude has not changed significantly over a predetermined period of time; the aircraft's speed is relatively constant; and the aircraft's throttle position has not changed. As shown at (204), method 200 includes adjusting the clearance (e.g., clearance 106) to a first clearance distance associated with the first flight condition. For example, in an embodiment, the clearance is adjusted (e.g., decreased) as the aircraft's speed increases.

[0062] As shown at (206), method 200 includes receiving, using the flight control system, a request for a second flight condition. For example, in an embodiment, the second flight condition may correspond to an increased power transient, such as a step climb power transient, caused by an aircraft operator initiating a step climb to a higher altitude. In some embodiments, the request may be received by monitoring the position of a throttle and observing a specific change in throttle angle, which indicates a step climb event. In another embodiment, an autopilot / autothrottle control system of flight control system 170 may also receive the request, such that the pilot requests an increase from a specific altitude to a second specific altitude. In response to the request, flight control system 170 sends a signal to engine control system 152 to request an increase in engine power.

[0063] Still refer to Figure 5 As shown in (208), method 200 includes adjusting at least two independently controllable parameters during the second flight condition, the at least two independently controllable parameters including at least a first parameter for optimizing clearance and a second parameter for satisfying the thrust requirement of the engine, the first parameter having a first effect on clearance, the second parameter having a second effect on clearance, and the first effect being greater than the second effect. For example, in an embodiment, the independently controllable parameters may include a power parameter of the engine (i.e., having a high effect on clearance) and a pitch parameter of the engine (i.e., having a lower effect on clearance) (e.g., regarding Figures 6A-6D ). More specifically, in embodiments, the power parameter may include fan speed, engine pressure ratio, or cooling flow. Additionally, in embodiments, the pitch parameter may include fan pitch. Thus, in embodiments, the method 200 may include independently controlling the power parameter and the pitch parameter, which may include, for example, simultaneously reducing the cooling flow in order to expand the shroud to increase clearance and prevent friction when the fan speed reaches high speeds.

[0064] Therefore, in certain embodiments, when the engine 10 is operated under tight operating clearance, adjusting the fan pitch and controlling the fan speed can minimize or otherwise reduce the transient turbine clearance effect without sacrificing thrust response. In other words, the fan speed is configured to minimize the clearance reduction, and the pitch adjustment is configured to control the thrust at a given fan speed. This requires independent control of the fan pitch and fan speed to meet the total fan thrust. In a further embodiment, the fan pitch can also be maintained below the aerodynamic and / or operability limits during this transient. In an additional embodiment, if multiple control handles are used and the gap closure can be reduced without sacrificing thrust response, other combinations in thrust control can also be used. Therefore, the method of the present disclosure provides a net thrust response to the second flight condition, which is essentially zero or negligible for the pilot. Further embodiments of the present disclosure can also be beneficial to the compressor operating clearance, because the overall rotor growth caused by the speed change can be balanced with the thermal growth of the rotor and stator components, thereby allowing the components and operating clearances to be tighter than traditional aircraft engines.

[0065] refer to Figures 6A-6D , the method 200 of the present disclosure can be better understood. Specifically, Fig. 6A is a graph 250 showing thrust as a function of time. In the illustrated embodiment, from time T1 to time T2, the engine is operated in a cruise condition 252. At time T2, the engine 10 receives a request for a second flight condition, such as a request for a step climb to a higher altitude. Due to the corresponding fan speed ( Figure 6B ) and fan pitch ( Fig.6D ) control, the thrust response is close to a normal thrust response, as shown by the close corresponding lines 254 (normal thrust response) and 256 (thrust response during the second flight condition).

[0066] Figure 6B 3 is a graph 300 showing a two-step engine response schedule as a function of fan speed. In the illustrated embodiment, dashed line 320 represents a simplified example of the engine response to command change 310 in the absence of clearance control system 150. Without the proposed two-step engine response schedule, after a request to increase fan speed from a steady-state cruising speed (represented as "0%" on the y-axis of graph 300) to a climb speed (e.g., an increase of approximately 12%), engine 10 responds quickly and increases the speed of the low speed rotor shaft to the climb speed.

[0067] The solid line represents a two-step response plan implemented using a gap control system 150 according to the present disclosure. Figure 6BIn the illustrated embodiment, the clearance control system 150 implements different engine response levels depending on the rate at which N1 increases from a steady-state cruise fan speed. Specifically, the clearance control system 150 implements a first "normal" level of engine response 330 ("fast response" 330) for smaller changes in fan speed, as represented by the larger slope during the fast response 330. In addition, as shown, the clearance control system 150 implements a second reduced level of engine response 340 ("slower response" 340) for larger changes in fan speed, as represented by the reduced slope during the slower response 340. In the illustrated embodiment, the clearance control system 150 implements the fast response 330 for changes in fan speed between 0% and about 4%, and implements the slower response 340 for changes in fan speed above 4%. In other words, the clearance control system 150 implements the fast response 330 for a first change in power level or a change in fan speed at a small increase in throttle, and implements the slower response 340 for a second change in power level or a change in fan speed at a larger increase in throttle.

[0068] The fast response 330 defines a rate limit for the increase of N1. In the illustrated embodiment, the fast response 330 limits the rate of increase of N1 to approximately 1.5% / second. Thus, the engine 10 responds to relatively small throttle increase requests with the fast response 330 for up to approximately 2-3 seconds. For rates of increase of N1 exceeding 4%, the slower response 340 defines a rate limit for the increase of N1 of approximately 0.25% / second, for example, until the fan speed increases by 12%. In other words, once N1 increases by 4% from its steady-state cruising speed, the clearance control system 150 implements the slower response 340 (slowing down the engine responsiveness) to increase N1 to the requested speed. Alternatively, for rates of increase of N1 that remain below 4%, the clearance control system 150 only implements the fast response 330.

[0069] This two-step response schedule facilitates maintaining fast engine response to relatively small throttle changes during routine maneuvers, for example, to maintain altitude and / or Mach number in response to environmental changes (e.g., wind gusts). In addition, the two-step response schedule that reduces engine responsiveness facilitates reducing the clearances 106 described herein, as well as the clearances between the turbine blades of the LP turbine 30 and the surrounding casing and other stages of the engine.

[0070] Reference now Fig.6D, a diagram 500 showing a two-step engine response plan as a function of fan pitch is shown. Specifically, as shown, the fan pitch can also be adjusted to achieve the command thrust under the second flight condition. For example, as shown, the fan pitch is maintained under steady-state cruise conditions. In addition, as shown, the command change in the flight condition is represented by line 510. Therefore, after receiving the demand for the second flight condition, the clearance control system 150 implements a normal pitch response in the first third or so of the second flight condition, which achieves interference suppression during level flight. Then, as shown at 506, the fan pitch can be increased more sharply to the operability / aerodynamic limit 508. Once such a limit 508 is reached, as shown at 512, the clearance control system 150 implements a slow change of the fan pitch close to the climb command, while the fan speed response keeps the total thrust constant. In addition, in such an embodiment, a tight operating clearance can be maintained during this period.

[0071] Special References Figure 6C , a graph 400 showing turbine clearance as a function of time is shown. Specifically, as shown, the change in clearance with the clearance control system 150 activated and not activated is represented by lines 402, 404. Therefore, as shown, the clearance control system 150 is configured to minimize clearance during the step climb transient, as shown by the change 406. Further, as shown, normal turbine clearance is represented by line 408, while reduced turbine clearance is represented by line 410. In other words, by using fan pitch to achieve commanded thrust and moving fan speed independently of fan pitch, the clearance control system 150 is configured to minimize the impact on turbine clearance.

[0072] Reference now Figure 7 , showing a method for controlling an aircraft engine (e.g., engine 10 (e.g., Figure 1 1) is a flow chart of an embodiment of a method 600 for determining a gap between a plurality of rotor blades and a shroud. Generally, the method 600 will be described herein with reference to the engine 10 described herein. However, it should be understood that the disclosed method 600 may be implemented with any engine having any other suitable configuration. In addition, although Figure 7 The steps performed in a particular order are depicted for the purpose of illustration and discussion, but the methods discussed herein are not limited to any particular order or arrangement. It will be understood by those skilled in the art using the disclosure provided herein that the various steps of the methods disclosed herein may be omitted, rearranged, combined, and / or adapted in various ways without departing from the scope of the present disclosure.

[0073] As shown in (602), method 600 includes receiving a request from an aircraft operator to step-climb an aircraft engine to a higher altitude using a flight control system. As shown in (604), method 600 includes increasing the engine speed of the aircraft engine to achieve the step-climb, while also adjusting the power parameters of the aircraft engine to optimize the clearance, and adjusting the fan pitch of the aircraft engine to meet the thrust demand of the aircraft engine. For example, in an embodiment, method 600 may include increasing the engine speed of the aircraft engine to achieve the step-climb, while also adjusting the power parameters, such as opening the pitch of the aircraft engine to meet the thrust demand of the step-climb. In addition, in an embodiment, when starting the step-climb, method 600 may include reducing the cooling flow to start the expansion of the shroud assembly 96, thereby increasing the turbine engine clearance to prevent friction when the fan speed is accelerated to the final high value. In such an embodiment, during the short duration step-climb transient, the larger transient clearance has the least effect on fuel combustion.

[0074] When it is determined that the aircraft has entered a stable flight phase and is expected to remain in the stable flight phase for a predetermined period of time, the systems, methods, and apparatus described herein have at least the technical effect of enabling the aircraft to operate more efficiently by reducing turbine clearance by a predetermined amount. Reducing turbine or compressor clearance helps improve fuel efficiency, thrust, and turbine life, each of which helps save fuel and / or service costs.

[0075] Exemplary embodiments of systems, methods, and apparatus for controlling turbine clearance in aircraft engines are described in detail above. The systems, methods, and apparatus are not limited to the specific embodiments described herein, but rather, the steps of the methods and / or components of the systems and / or apparatus may be used independently of other steps and / or components described herein. In addition, the steps and / or components may also be defined in other systems, methods, and / or apparatus, or used in conjunction therewith, and are not limited to being practiced using only the systems, methods, and apparatus described herein. Other embodiments may include actuation of an existing engine variable geometry, which may include variable stator vanes, variable stator inlet guide vanes, variable exhaust valves, customer or domestic exhaust valves, regulating turbine cooling systems, and / or third flow regulating doors.

[0076] Further aspects are provided by the subject matter of the following clauses:

[0077] A method for reducing a gap between a plurality of rotor blades and a shroud assembly of an engine, the method comprising: determining, using a flight control system, that an aircraft is in a first flight condition; adjusting the gap to a first gap distance associated with the first flight condition; receiving, using the flight control system, a requirement for a second flight condition; and during the second flight condition, adjusting at least two independently controllable parameters, the at least two independently controllable parameters including at least a first parameter for optimizing the gap and a second parameter for meeting a thrust requirement of the engine, the first parameter having a first effect on the gap, the second parameter having a second effect on the gap, the first effect being greater than the second effect.

[0078] A method as in any preceding clause, wherein the at least two independently controllable parameters comprise a power parameter of the engine and a pitch parameter of the engine.

[0079] A method as in any preceding clause, wherein the power parameter comprises at least one of fan speed, engine pressure ratio or cooling flow, and the pitch parameter comprises fan pitch.

[0080] A method as in any preceding clause, further comprising maintaining the fan pitch below aerodynamic and operability limitations.

[0081] A method as in any preceding clause, wherein the first flight condition is a steady state cruise condition.

[0082] A method as in any preceding clause, wherein the second flight condition is an increased power transient.

[0083] A method as in any preceding clause, wherein a net thrust response to the second flight condition is negligible.

[0084] A method as in any preceding clause, wherein the engine comprises at least one of a hybrid electric engine, an electric engine or a fuel combustion engine.

[0085] A system for reducing a gap between a plurality of rotor blades and a shroud assembly of an engine, the system comprising: a flight control system comprising a memory and one or more processors, the one or more processors being configured to perform a plurality of operations, the plurality of operations comprising: determining that an aircraft is in a first flight condition; adjusting the gap to a first gap distance associated with the first flight condition; receiving a demand for a second flight condition; and during the second flight condition, adjusting at least two independently controllable parameters, the at least two independently controllable parameters comprising at least a first parameter for optimizing the gap and a second parameter for meeting a thrust demand of the engine, the first parameter having a first effect on the gap, the second parameter having a second effect on the gap, the first effect being greater than the second effect.

[0086] A system as claimed in any preceding clause, wherein the at least two independently controllable parameters comprise a power parameter of the engine and a pitch parameter of the engine.

[0087] A system as claimed in any preceding clause, wherein the power parameter comprises at least one of fan speed, engine pressure ratio or cooling flow, and the pitch parameter comprises fan pitch.

[0088] The system of any preceding clause, wherein the plurality of operations further comprises maintaining the fan pitch below aerodynamic and operability limitations.

[0089] A system as described in any preceding clause, wherein the first flight condition is a steady state cruise condition.

[0090] A system as in any preceding clause, wherein the second flight condition is an increased power transient.

[0091] A system as in any preceding clause, wherein the engine comprises at least one of a hybrid electric engine, an electric engine, or a fuel combustion engine.

[0092] A method for controlling clearance between a plurality of rotor blades and a shroud assembly of an aircraft engine, the method comprising: receiving, using a flight control system, a request from an aircraft operator to step climb the aircraft engine to a higher altitude; and increasing an engine speed of the aircraft engine to achieve the step climb while also adjusting a power parameter of the aircraft engine to optimize the clearance and adjusting a fan pitch of the aircraft engine to meet a thrust demand of the aircraft engine.

[0093] A method as in any preceding clause, wherein the power parameter comprises at least one of fan speed, engine pressure ratio or cooling flow.

[0094] A method as in any preceding clause, further comprising independently controlling said power parameter and said fan pitch.

[0095] A method as in any preceding clause, further comprising maintaining the fan pitch below aerodynamic and operability limitations.

[0096] A method as in any preceding clause, wherein the aircraft engine comprises at least one of a hybrid electric engine, an electric engine or a fuel combustion engine.

[0097] This written description uses exemplary embodiments to describe the presently disclosed subject matter, including the best mode, and also to enable those skilled in the art to practice such subject matter, including making and using any devices or systems and performing any combined methods. The patentable scope of the presently disclosed subject matter is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or include equivalent structural elements that do not differ substantially from the literal language of the claims.

Claims

1. A method for reducing a gap between a plurality of rotor blades and a shroud assembly of an engine, characterized in that: The method comprises: determining, using a flight control system, that the aircraft is in a first flight condition; adjusting the gap to a first gap distance associated with the first flight condition; receiving, with the flight control system, a request for a second flight condition; and During the second flight condition, at least two independently controllable parameters are adjusted, the at least two independently controllable parameters including at least a first parameter for optimizing the gap and a second parameter for meeting the thrust requirement of the engine, the first parameter having a first effect on the gap, the second parameter having a second effect on the gap, the first effect being greater than the second effect.

2. The method according to claim 1, characterized in that: The at least two independently controllable parameters include a power parameter of the engine and a pitch parameter of the engine.

3. The method according to claim 2, characterized in that Wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio or a cooling flow rate, and the pitch parameter comprises a fan pitch.

4. The method according to claim 3, characterized in that Further included maintaining the fan pitch below aerodynamic and operability limitations.

5. The method according to claim 1, characterized in that The first flight condition is a steady-state cruising condition.

6. The method according to claim 1, characterized in that The second flight condition is a power increase transient.

7. The method according to claim 1, characterized in that The net thrust response to the second flight condition is negligible.

8. The method according to claim 1, characterized in that The engine comprises at least one of a hybrid electric engine, an electric engine or a fuel combustion engine.

9. A system for reducing the gap between a plurality of rotor blades and a shroud assembly of an engine, characterized in that: The system comprises: A flight control system comprising a memory and one or more processors, wherein the one or more processors are configured to perform a plurality of operations, the plurality of operations comprising: determining that the aircraft is in a first flight condition; adjusting the gap to a first gap distance associated with the first flight condition; receiving a request for a second flight condition; and During the second flight condition, at least two independently controllable parameters are adjusted, the at least two independently controllable parameters including at least a first parameter for optimizing the gap and a second parameter for meeting the thrust requirement of the engine, the first parameter having a first effect on the gap, the second parameter having a second effect on the gap, the first effect being greater than the second effect.

10. The system according to claim 9, characterized in that The at least two independently controllable parameters include a power parameter of the engine and a pitch parameter of the engine.

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