Systems and methods for reducing clearances in engines

Abstract

A method for reducing the clearance between multiple rotor blades and a shroud assembly of an engine includes: determining, using a flight control system, that the aircraft is in a first flight condition. The method further includes adjusting the clearance to a first clearance distance associated with the first flight condition. Additionally, the method includes receiving a demand for a second flight condition using the flight control system. During the second flight condition, the method includes adjusting at least two independently controllable parameters, the at least two independently controllable parameters including at least a first parameter for optimizing the clearance and a second parameter for meeting the engine's thrust requirements, the first parameter having a first effect on the clearance and the second parameter having a second effect on the clearance, the first effect being greater than the second effect.

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to aircraft engines, and more particularly, to controlling clearances within an aircraft engine to facilitate more efficient operation of the aircraft engine during operation. BACKGROUND

[0002] At least some known aircraft include an engine control system, sometimes referred to as a full authority digital engine control (FADEC). The FADEC is a system that includes a digital computer and its associated accessories for controlling various aspects of aircraft engine performance. The FADEC receives a plurality of current input variables of 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 these data, engine operating parameters such as fuel flow, stator vane position, exhaust valve position, etc. are calculated and applied as appropriate to provide optimum engine efficiency under the given current flight conditions.

[0003] Aircraft also typically include a flight control system, which can include a system commonly referred to as a flight management system (FMS). The FMS is a specialized computer system that can automatically perform various flight tasks, including in-flight management of a flight plan. Using various sensors, such as but not limited to a global positioning system (GPS), an inertial navigation system (INS), and supplemented with radio navigation to determine the position of the aircraft, 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 touchscreen. The FMS transmits the flight plan for display on an EFIS, a navigation display (ND), or a multifunction display (MFD). The FADEC and the FMS are independent systems that can communicate current values of parameters in certain circumstances.

[0004] Some known aircraft engines include hot section modules and cold section modules that include compressors and turbines. To improve fuel efficiency, thrust, and / or turbine life, at least some known engines attempt to control the distance or clearance 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, can increase. Both of these effects can be due to increased fuel flow during maneuvers such as climb, descent / landing sequences, and / or evasive maneuvers. Additionally, during operation, particularly during transient operations such as those requiring increased fuel flow, the blade tip length can expand faster than the shroud or stator. Thus, during such operations, the blade tip can come into contact with the shroud, a condition referred to as rub.

[0005] At least some known aircraft engines use active clearance control to prevent rub while minimizing operating clearances under all operating conditions. In at least some conventional systems, active clearance control attempts to expand or contract shrouds or stators 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 hardware mechanical growth due to changes in speed.

[0006] Further, at least some known aircraft engines activate clearance control in response to one or more engine operating parameters. Further, at least some known aircraft engines activate clearance control based on elapsed time relative to transient engine conditions, such as throttle manipulation and / or rotor speed changes. Further, 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 can not adequately predict increases in fuel flow to mitigate rub of blade tips due to clearance being too small. For example, during a flight that requires a change in throttle to climb from one altitude to another, the aircraft engine response is typically based on a predetermined schedule, which results in rotor blades growing (e.g., lengthening) faster than the shrouds surrounding them due to mechanical acceleration of the rotor blades. As the engine speed increases, the clearance control system lags behind the relatively fast expansion of the blades. Accordingly, there is a need in the art for improved control strategies to prevent rub. BRIEF DESCRIPTION OF DRAWINGS

[0008] A complete and enabling disclosure is set forth in the specification of this patent document, including the best mode for carrying out the application, of which the exemplary aspects are a part, and which is made with reference to the drawings, wherein:

[0009] Figure 1 is a cross-sectional view of a gas turbine engine in accordance with exemplary aspects of the present disclosure.

[0010] Figure 2 is a cross-sectional view of a gas turbine engine in accordance with exemplary aspects of the present disclosure. Figure 1 is an enlarged cross-sectional schematic view of a portion of a gas turbine engine including a clearance control system.

[0011] Figure 3 is an enlarged cross-sectional schematic view of a portion of a gas turbine engine including a clearance control system. Figure 2 is an enlarged cross-sectional schematic view of a portion of a gas turbine engine including a clearance control system.

[0012] Figure 4 is a schematic block diagram of a system for reducing turbine or compressor clearances between a plurality of rotor blades of a turbine engine and a shroud of the turbine engine in accordance with exemplary embodiments of the present application.

[0013] Figure 5is a flowchart of a method for reducing turbine clearances between a plurality of rotor blades of a turbine engine and a shroud of the turbine engine according to an example embodiment of the present disclosure.

[0014] Figure 6A is a chart showing a comparison of a current fan speed response of an aircraft engine to a fan speed of an aircraft engine using a clearance control system according to an embodiment of the present disclosure.

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

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

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

[0018] Figure 7 is a flowchart 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] Repetition of reference characters in the specification and drawings is intended to indicate the same or similar features or elements. 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 drawings. Each example is provided by way of explanation of the present disclosure and is not meant as a limitation of the present disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present disclosure without departing from the scope or spirit of the present disclosure. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield still a further embodiment. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.

[0021] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless otherwise explicitly provided, none of the embodiments described herein are intended to be exclusive of any other embodiments. In fact, as would be recognized by one skilled in the art, any of the embodiments described herein can be combined with one or more of the other embodiments described herein.

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

[0023] For example, the term “at least one of A, B, and C” in the context 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-producing section (e.g., a combustion section), and one or more turbines together to produce a torque output.

[0025] The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, and the like, as well as hybrid-electric versions of one or more of these engines.

[0026] The term “combustion section” refers to any heat adding system of a turbomachine. For example, the term “combustion section” can 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 adding assembly. In certain example embodiments, a combustion section can include an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion system, or combinations thereof.

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

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

[0029] The terms “low” and “high” or their respective comparative forms (e.g., lower, higher, as applicable) when used in conjunction with a compressor, turbine, shaft or spool component, or the like, both refer to relative speeds within an engine, unless otherwise noted. For example, a “low turbine” or “low speed turbine” defines a component that is configured to operate at a lower rotational speed (e.g., maximum allowable rotational speed) than a “high turbine” or “high speed turbine” of an engine.

[0030] The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle and refer to a normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, forward refers to a position closer to an engine inlet, and aft refers to a position closer to an engine nozzle or exhaust.

[0031] The terms“upstream” and“downstream” refer to the relative direction with respect to the 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.

[0032] As used herein, the terms“axial” and“axially” refer to a direction and an orientation that extends substantially parallel to a centerline of the gas turbine engine. Further, the terms“radial” and“radially” refer to a direction and an orientation that extends substantially perpendicular to the centerline of the gas turbine engine. Further, as used herein, the terms“circumferential” and“circumferentially” refer to a direction and an orientation that extends arcuately about the centerline of the gas turbine engine.

[0033] The terms“coupled,”“fixed,”“attached to,” and the like, mean either a direct coupling, fixation, or attachment, or an indirect coupling, fixation or attachment via one or more intermediary components or features, unless specifically stated to the contrary.

[0034] As used herein, the terms“first,”“second,”“third,” etc. can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

[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 that the two walls and / or surfaces are in a serial contact relationship with the one or more non-structural layers (i.e., the 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“one-piece” with respect to a structure means a structure that is formed from a continuous material or group of materials, without seams, connecting joints, etc. The integral, one-piece structures described herein can be formed by additive manufacturing to have the described structure, or by a casting process, etc.

[0037] Approximating language as used herein with respect to a given value or range of values should generally be construed to encompass not only the value or range of values itself but also an amount that is functionally or structurally close thereto as would be recognized by one of ordinary skill in the art. For example, a value recited as“about X” should generally be construed to include not only the value of X but also a range of values up to and including X — for example, an amount that performs a function substantially the same as X or is structurally or functionally similar to X. Unless otherwise specified, all measurements are understood to be made at standard conditions, e.g., about 25 °C and 1 atmosphere pressure. Approximating language can be used to connote the approximate but not exact nature of one or more parameters. For example, sometimes one or more parameters encompassed by a given value or range of values are approximated, e.g., as“about X,”“generally X,” or“substantially X.” It will be understood by those skilled in the art that such parameters, which can take on values approximately the same as the stated values, can not be limited to the same precision as the value itself. For example, a lower precision of measurement or a less precise value resulting from a rule of thumb, estimation, or approximation is encompassed by a given value or range of values. It will be further understood that the values are approximations which are now more precise or amenable to more precise measurements.

[0038] Herein and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges within those ranges except where it would be obvious, such as might be stated in the prior art. For example, a range of "about 1 w / w-% to about 5 w / w-%" is interpreted to include from about 1 w / w-% to about 5 w / w-% not inclusive of the endpoints.

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

[0040] Reference will now be made to the drawings wherein like numerals refer to like components throughout. Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 in accordance with an example embodiment of the present disclosure. More specifically, for Figure 1 Embodiments of the gas turbine engine 10 are high-bypass turbofan engines, sometimes referred to as "turbofan engines." As shown schematically in Figure 1 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 section 16 located downstream from the fan section 14.

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

[0042] In the illustrated embodiment, fan section 14 includes a fan 38 having a plurality of fan blades 40 spaced apart and coupled to a disk 42. As shown, the fan blades 40 extend outward from the disk 42 generally in a radial direction R. Each fan blade 40 is operably coupled to a pitch changing mechanism 44 by means of the fan blades 40, the pitch changing mechanism 44 being rotatable relative to the disk 42 about a pitch axis P, the pitch changing mechanism 44 being configured to collectively change the pitch of the fan blades 40, for example, uniformly changing the pitch of the fan blades 40. The gas turbine engine 10 may also include a power gearbox 46, the fan blades 40, the disk 42, and the pitch changing mechanism 44 rotating together about a longitudinal centerline 12 via a 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 LP shaft 36, allowing the fan 38 to rotate at a more efficient fan speed. Therefore, the gas turbine engine 10 shown is configured as a geared gas turbine engine (i.e., including a power gearbox 46) and a variable pitch gas turbine engine (i.e., including a fan 38 configured as a variable pitch fan).

[0043] Still referencing Figure 1 In this embodiment, the disc 42 is covered by a rotatable front hub 48 (sometimes referred to as a "rotor") of the fan section 14. The front hub 48 has an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. Furthermore, the fan section 14 includes an annular fan housing or outer nacelle 50 that circumferentially surrounds at least a portion of the fan 38 and / or 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. Additionally, a downstream section 54 of the nacelle 50 extends over the outer portion of the turbine 16, thereby defining a bypass airflow passage 56 between them.

[0044] During operation of the gas turbine engine 10, a certain amount of air 58 enters the gas turbine engine 10 through the nacelle 50 and the corresponding inlets 60 of the fan section 14. As this certain amount of air 58 passes through the fan blades 40, a first portion of air 62 is directed or diverted into the bypass airflow passage 56, while a second portion of air 64 is directed or diverted into the working gas flow path 37, as indicated by arrow 64, or more specifically, into the LP compressor 22. The ratio between the first portion of air 62 and the second portion of air 64 is commonly referred to as the 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, whereby the second portion of air 64 mixes with fuel and burns 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 vanes 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 vanes 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 ejection 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 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, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the ejection 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 illustrated gas turbine engine 10 is by way of example only, and in other example embodiments, the gas turbine engine 10 can have any other suitable configuration. For example, although the illustrated gas turbine engine 10 is configured as a ducted gas turbine engine (i.e., includes the outer nacelle 50), in other embodiments, the gas turbine engine 10 can 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 aspects of the present disclosure can be incorporated into any other suitable gas turbine engine in other embodiments. For example, in other embodiments, aspects of the present disclosure can be incorporated into, as appropriate, a turboprop gas turbine engine, a turboshaft gas turbine engine, a turbojet gas turbine engine, or a hybrid-electric engine. In additional embodiments, aspects of the present disclosure can be incorporated into engines that burn hydrogen gas, sustainable aviation fuel, or are driven by hydrogen fuel cells.

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

[0049] like Figure 3 As shown, the annular shroud assembly 96, also referred to as the static housing assembly, surrounds a row of HP turbine rotor blades 70 and maintains a tight clearance relationship with them. In an embodiment, the shroud assembly 96 extends radially inward from the surrounding turbine housing 98 and includes a plurality of shroud members 100 ( Figure 3 (or arc-shaped sector.) In addition, as shown in the figure, adjacent shield members 100 are connected together such that the shield member 100 surrounds the HP turbine rotor blade 70.

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

[0051] Therefore, the clearance control system 150 helps to control the clearance 106 during engine operation. Therefore, now referring to... Figure 4 A schematic block diagram of an embodiment of a gap control system 150 according to the present disclosure is shown. In an embodiment, the gap 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 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 related to the engine 154. These process parameters can include, for example, engine speed, fuel flow, damper and guide vane positions, stator vane clearances, 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 can 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. Further, as shown, the flight control system 170 includes a processor 174 and a memory 176 communicatively coupled to the processor 174. In embodiments, 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 can be a wireless communication medium. In embodiments, the flight control system 170 is located proximate to a cockpit (not shown) of the aircraft, and the controller 152 is located proximate to the engine with which it is associated. The flight control system 170 can be embodied in a single processor-based component, or the functions of the flight control system 170 can be performed by a plurality of components configured to perform the functions described herein. Some components performing the functions of the flight control system 170 can be located proximate to the cockpit, while other components can be distributed throughout the aircraft for ease, 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 communication between the engine controller and any avionics functionality 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 can receive current aircraft status from a plurality of aircraft sensors 178 through a sensing system 180. Such aircraft sensors 178 can include pitot tubes for determining airspeed, gyroscopes, compasses, accelerometers, position sensors, altimeters, and various other sensors capable of detecting a condition, state, or position of the aircraft.

[0055] Further, as Figure 4 shown, the flight control system 170 can also receive information from one or more onboard processing systems 182, which can be standalone systems or systems having functionality distributed across multiple computer systems. Further, the flight control system 170 and the onboard processing systems 182 can use wired communication channels and / or network connections (e.g., Ethernet or fiber optic), wireless communication means (e.g., radio frequency (RF), such as FM radio and / or digital audio broadcast, Institute of Electrical and Electronics Engineers 802.11 standards (e.g., 802.11(g) or 802.11(n)), global system for mobile communications (GSM), satellite communication links, etc. and / or any other suitable communication means.

[0056] As used herein, a wired communication pathway includes pathways that use optical fibers and other optical devices for communication. The flight control system 170 can also receive information from one or more off-board processing systems 184, which can be stand-alone systems or systems having functionality distributed across multiple computer systems and / or multiple sites. The off-board processing systems 184 and the 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 broadcast, Institute of Electrical and Electronics Engineers 802.11 standards (e.g., 802.11(g) or 802.11(n)), global system for mobile communications (GSM), satellite communication links, etc. and / or any other suitable communication means.

[0057] As with at least some known aircraft operating procedures, a step climb maneuver occurs when an aircraft pilot elects to increase the altitude of the aircraft flight. According to current FAA regulations, altitude steps are typically incremented by 2,000 feet. This means, for example, that an aircraft pilot flying at an altitude of 33,000 feet can elect to perform a step climb maneuver to cause the aircraft to climb to an altitude of 35,000 feet. To implement the step climb maneuver, the pilot modifies the controls of the autopilot / autothrottle system of the flight control system 170 to request the aircraft to ascend to the desired cruise altitude. The flight control system 170 then uses a predetermined algorithm to increase engine power to cause the aircraft to climb. Since the requirement to increase engine power typically requires the gas turbine engine 10 to spin faster to increase engine thrust, the HP turbine rotor blades 70 grow due to mechanical forces and associated thermal changes. This turbine blade growth can cause clearances within, for example, the HP turbine 28 to decrease. If this growth exceeds the designed clearances, the HP turbine rotor blades 70 will rub against the turbine casing 98 of the gas turbine engine 10, potentially damaging engine components or reducing engine efficiency.

[0058] For example, an aircraft step climb from approximately 33,000 feet to 35,000 feet can 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. Because the HP turbine rotor blades 70 grow faster than the surrounding turbine casing 98, engine designers must account for additional clearance to prevent rubbing under these conditions. In the above example, the additional clearance is referred to herein as a step climb clearance. However, by increasing the clearance between the ends of the HP turbine rotor blades 70 and the turbine casing 98, more air is able to escape through the HP turbine rotor blades 70 rather than threading through the HP turbine rotor blades 70, which results in decreased engine performance and increased fuel consumption. Therefore, there is a need to develop algorithms that account for 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 shrink the turbine shroud to reduce the clearance in cruise mode, and reducing the cooling flow to increase the clearance during step climb so that rubbing does not occur when the fan speed reaches high values. Due to the thermal time constant of the turbine shroud, the clearance does not instantaneously increase, but rather gradually increases so that when the fan speed accelerates to high values the clearance is large enough to prevent rubbing.

[0060] Referring now to Figure 5 , a flow diagram of an embodiment of a method 200 of reducing a clearance between a plurality of rotor blades and a shroud of an engine, such as the engine 10 (as shown in Figure 1 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 can be implemented with any engine having any other suitable configuration. Moreover, although the steps of the method 200 are depicted in a particular order for illustrative and discussion purposes, the methods discussed herein are not limited to any particular order or arrangement. Those skilled in the art, using the disclosures provided herein, will understand that the various steps of the methods disclosed herein can be omitted, rearranged, combined, and / or adapted in various ways without deviating from the scope of the present disclosure. Figure 5 For illustrative and discussion purposes, steps are depicted as being performed in a particular order. However, the methods discussed herein are not limited to any particular order or arrangement. Those skilled in the art, using the disclosures provided herein, will understand that the various steps of the methods disclosed herein can be omitted, rearranged, combined, and / or adapted in various ways without deviating from the scope of the present disclosure.

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

[0062] As shown at (206), the method 200 includes receiving, with the flight control system, a demand for a second flight condition. For example, in an embodiment, the second flight condition can correspond to an increase power transient due to the aircraft operator beginning a step climb to a higher altitude, e.g., a step climb power transient. In certain embodiments, the demand can be received by monitoring the position of the throttle and observing a particular change in the throttle angle, which indicates a step climb event. In another embodiment, the autopilot / automatic throttle control system of the flight control system 170 can also receive the request, such that the pilot requests an increase from one particular altitude to a second particular altitude. In response to the request, the flight control system 170 sends a signal to the engine control system 152 to request an increase in engine power.

[0063] Still referring to Figure 5 As shown at (208), the method 200 includes adjusting, during the second flight condition, at least two independently controllable parameters, the at least two independently controllable parameters including at least a first parameter for optimizing the clearance and a second parameter for satisfying a thrust demand of the engine, the first parameter having a first effect on the clearance and the second parameter having a second effect on the clearance, and the first effect being greater than the second effect. For example, in an embodiment, the independently controllable parameters can include a power parameter of the engine (i.e., having a high effect on the clearance) and a pitch parameter of the engine (i.e., having a lower effect on the clearance) (as described with respect to FIGS. 2A-2B, for example). More specifically, in an embodiment, the power parameter can include a fan speed, an engine pressure ratio, or a cooling flow. Further, in an embodiment, the pitch parameter can include a fan pitch. Accordingly, in an embodiment, the method 200 can include independently controlling the power parameter and the pitch parameter, which can include, for example, simultaneously decreasing the cooling flow in order to expand the shroud to increase the clearance, preventing rubs from occurring when the fan speed reaches high speeds. Figures 6A-6D Further shown and described). More specifically, in an embodiment, the power parameter can include a fan speed, an engine pressure ratio, or a cooling flow. Further, in an embodiment, the pitch parameter can include a fan pitch. Accordingly, in an embodiment, the method 200 can include independently controlling the power parameter and the pitch parameter, which can include, for example, simultaneously decreasing the cooling flow in order to expand the shroud to increase the clearance, preventing rubs from occurring when the fan speed reaches high speeds.

[0064] Accordingly, in certain embodiments, when engine 10 is operating at tight operating clearances, adjusting fan pitch and controlling fan speed can minimize or otherwise reduce the transient turbine clearance impact without sacrificing thrust response. In other words, fan speed is configured to minimize clearance reduction, while pitch adjustment is configured to control thrust at a given fan speed. This requires independent control of fan pitch and fan speed to meet total fan thrust. In further embodiments, fan pitch can also be maintained below aeromechanics and / or operability limits during this transient. In additional embodiments, other combinations in thrust control can also be used if multiple control handles are used and clearance closure can be reduced without sacrificing thrust response. Accordingly, the methods of the present disclosure provide a net thrust response to the second flight condition that is substantially zero or negligible to the pilot. Further embodiments of the present disclosure can also benefit compressor operating clearances, as overall rotor growth due to speed changes can be balanced with thermal growth of rotor and stator components, allowing for tighter component and operating clearances than conventional aircraft engines.

[0065] Reference is made to Figures 6A-6D The methods of the present disclosure can be better understood with reference to FIG. 2. In particular, Figure 6A is a graph 250 showing thrust as a function of time. In the illustrated embodiment, from time Tl to time T2, the engine is operating at cruise conditions 252. At time T2, engine 10 receives a demand for a second flight condition, such as a demand for a step climb to a higher altitude. Due to the corresponding fan speed Figure 6B ) and fan pitch Figure 6D ) controls, the thrust response is near normal thrust response, as shown by the nearness of the corresponding lines 254 (normal thrust response) and 256 (thrust response during the second flight condition).

[0066] Figure 6B is a graph 300 showing a two-step engine response schedule as a function of fan speed. In the illustrated embodiment, the dashed line 320 represents a simplified example of engine response to a command change 310 in the absence of a clearance control system 150. In the absence of the proposed two-step engine response schedule, after a request to increase fan speed from a steady-state cruise speed (represented as "0%" on the y-axis of graph 300) to a climb speed (e.g., an increase of about 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 schedule implemented using a clearance control system 150 according to the present disclosure. In the illustrated embodiment, the response schedule is implemented by adjusting fan pitch and controlling fan speed to maintain a constant total fan thrust. In the illustrated embodiment, the response schedule is implemented by adjusting fan pitch and controlling fan speed to maintain a constant total fan thrust. Figure 6BIn the illustrated embodiment, the clearance control system 150 implements different levels of engine responsiveness depending on the rate of increase of Nl from the steady state cruise fan speed. Specifically, the clearance control system 150 implements a first, "normal" level of engine responsiveness 330 ("fast response" 330) for small changes in fan speed, as represented by the larger slope during the fast response 330. In addition, as illustrated, the clearance control system 150 implements a second, reduced level of engine responsiveness 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 changes in fan speed for a first change in power level or small increase in throttle, and implements the slower response 340 for changes in fan speed for a second change in power level or large increase in throttle.

[0068] The fast response 330 defines a limit on the rate of increase of Nl. In the illustrated embodiment, the fast response 330 limits the rate of increase of Nl to about 1.5% / sec. Thus, the engine 10 responds to a relatively small throttle increase request with the fast response 330 for a duration of up to about 2-3 seconds. For increases in Nl above 4%, the slower response 340 defines a limit on the rate of increase of Nl of about 0.25% / sec, for example, up to a 12% increase in fan speed. In other words, once Nl has increased 4% from its steady state cruise speed, the clearance control system 150 implements the slower response 340 (slowing the engine responsiveness) to increase Nl to the requested speed. Alternatively, for Nl increase rates that remain below 4%, the clearance control system 150 implements only the fast response 330.

[0069] This two-step response plan is advantageous for maintaining a fast engine response for 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 plan for reducing engine responsiveness is advantageous for reducing the clearances 106 described herein, as well as clearances between turbine blades of the LP turbine 30 and the surrounding casing and between stages of the engine.

[0070] Reference is now made to Figure 6DA graph 500 illustrates a two-step engine response plan as a function of fan pitch. Specifically, as shown, the fan pitch can also be adjusted under a second flight condition to achieve the commanded thrust. For example, as shown, the fan pitch is maintained under steady-state cruise conditions. Furthermore, as shown, the commanded changes in flight conditions are represented by line 510. Thus, upon receiving a demand for the second flight condition, the clearance control system 150 implements a normal pitch response for approximately one-third of the second flight condition, achieving disturbance suppression during level flight. Then, as shown in 506, the fan pitch can be increased more abruptly to the operability / aerodynamic limit 508. Once such a limit 508 is reached, as shown in 512, the clearance control system 150 implements a slow change in fan pitch approaching the climb command, while the fan speed response maintains a constant total thrust. Furthermore, in such an embodiment, a tight operating clearance can be maintained during this period.

[0071] Special Reference Figure 6C A graph 400 is shown illustrating turbine clearance as a function of time. Specifically, as shown, the change in clearance when the clearance control system 150 is activated and deactivated is represented by lines 402 and 404. Therefore, as shown, the clearance control system 150 is configured to minimize clearance during the step-climb transient, as shown in variation 406. Furthermore, 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 effect on turbine clearance.

[0072] Now for reference Figure 7 This shows the control of the aircraft engine (e.g., engine 10). Figure 1 The flowchart illustrates an embodiment of a method 600 for clearing the gap between multiple rotor blades and their shrouds. Generally, method 600 will be described herein with reference to the engine 10 described herein. However, it should be understood that the disclosed method 600 can be implemented with any engine having any other suitable configuration. Furthermore, although... Figure 7 The steps performed in a particular order are depicted for illustrative and discussion purposes, but the methods discussed herein are not limited to any particular order or arrangement. Those skilled in the art will understand using the disclosure provided herein that the steps of the methods disclosed herein may be omitted, rearranged, combined, and / or adapted in various ways without departing from the scope of this disclosure.

[0073] As shown at (602), the method 600 includes receiving, with the flight control system, a request from a flight operator to step climb the aircraft engine to a higher altitude. As shown at (604), the method 600 includes 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 clearance, and adjusting a fan pitch of the aircraft engine to meet thrust requirements of the aircraft engine. For example, in an embodiment, the method 600 can include increasing the engine speed of the aircraft engine to achieve the step climb while also adjusting the power parameter, such as opening the pitch of the aircraft engine to meet the thrust requirements of the step climb. Further, in an embodiment, upon starting the step climb, the method 600 can include reducing the cooling flow to start expansion of the shroud assembly 96, thereby increasing the turbine engine clearance to prevent rubbing as the fan speed accelerates to the final high value. In such an embodiment, the larger transient clearance has minimal impact on fuel burn during the short duration step climb transient.

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

[0075] The exemplary embodiments of systems, methods, and apparatus for controlling turbine clearance in an aircraft engine 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 the components of the systems and / or apparatus can be utilized independently of other steps and / or components described herein. Further, the steps and / or components described can also be defined in, or used in combination with, other systems, methods, and / or apparatus, and are not limited to practice only with the systems, methods, and apparatus as described herein. Other embodiments can include actuation of existing engine variable geometry, which can include variable stator vanes, variable stator inlet guide vanes, variable bleed valves, customer or home bleed valves, regulated turbine cooling systems, and / or third flow regulation doors.

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

[0077] A method for reducing a clearance between a plurality of rotor blades and a shroud assembly of an engine, the method comprising: determining, with a flight control system, that an aircraft is in a first flight condition; adjusting the clearance to a first clearance distance associated with the first flight condition; receiving, with the flight control system, 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 clearance and a second parameter for meeting a thrust demand of the engine, the first parameter having a first effect on the clearance, the second parameter having a second effect on the clearance, the first effect being greater than the second effect.

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

[0079] The method of any preceding paragraph, wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio, or a cooling flow, and the pitch parameter comprises a fan pitch.

[0080] The method of any preceding paragraph, further comprising maintaining the fan pitch below aeromechanics and operability limits.

[0081] The method of any preceding paragraph, wherein the first flight condition is a steady state cruise condition.

[0082] The method of any preceding paragraph, wherein the second flight condition is an increase power transient.

[0083] The method of any preceding paragraph, wherein a net thrust response to the second flight condition is negligible.

[0084] The method of any preceding paragraph, wherein the engine comprises at least one of a hybrid-electric engine, an electric engine, or a fuel burning engine.

[0085] A system for reducing a clearance 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 configured to perform a plurality of operations comprising: determining that an aircraft is in a first flight condition; adjusting the clearance to a first clearance 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 comprising at least a first parameter for optimizing the clearance and a second parameter for meeting a thrust demand of the engine, the first parameter having a first effect on the clearance, the second parameter having a second effect on the clearance, the first effect being greater than the second effect.

[0086] The system of 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] The system of any preceding clause, wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio, or a cooling flow, and the pitch parameter comprises a fan pitch.

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

[0089] The system of any preceding clause, wherein the first flight condition is a steady state cruise condition.

[0090] The system of any preceding clause, wherein the second flight condition is an increase power transient.

[0091] The system of any preceding clause, wherein the engine comprises at least one of a hybrid-electric engine, an electric engine, or a fuel burning engine.

[0092] A method for controlling a clearance between a plurality of rotor blades and a shroud assembly of an aircraft engine, the method comprising: receiving, with 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] The method of any preceding clause, wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio, or a cooling flow.

[0094] The method of any preceding clause, further comprising independently controlling the power parameter and the fan pitch.

[0095] The method of any preceding clause, further comprising maintaining the fan pitch below aerodynamic and operability limits.

[0096] The method of any preceding clause, wherein the aircraft engine comprises at least one of a hybrid-electric engine, an electric engine, or a fuel-burning engine.

[0097] This written description uses examples to describe the presently disclosed subject matter, including the best mode, and also to enable any person skilled in the art to practice the same, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the presently disclosed subject matter is defined by the claims, and can 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 if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Claims

1. A method for reducing the clearance between a plurality of rotor blades and a shroud assembly of an engine, characterized by, The method comprises: determining, with 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, with the flight control system, 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, wherein a net thrust response to the second flight condition is substantially zero.

2. The method of claim 1, wherein, wherein the at least two independently controllable parameters comprise a power parameter of the engine and a pitch parameter of the engine.

3. The method of claim 2, wherein, wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio, or a cooling flow, and the pitch parameter comprises a fan pitch.

4. The method of claim 3, wherein, further comprising maintaining the fan pitch below aerodynamic and operability limits.

5. The method of claim 1, wherein, wherein the first flight condition is a steady state cruise condition.

6. The method of claim 1, wherein, wherein the second flight condition is an increase power transient.

7. The method of claim 1, wherein, wherein the engine comprises at least one of a hybrid electric engine, an electric engine, or a fuel burning engine.

8. A system for reducing the clearance between a plurality of rotor blades and a shroud assembly of an engine, characterized by, The system comprises: a flight control system comprising a memory and one or more processors 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, wherein a net thrust response to the second flight condition is substantially zero.

9. The system of claim 8, wherein, wherein the at least two independently controllable parameters comprise a power parameter of the engine and a pitch parameter of the engine.

10. The system of claim 9, wherein, wherein the power parameter comprises at least one of a fan speed, an engine pressure ratio, or a cooling flow, and the pitch parameter comprises a fan pitch.

11. The system of claim 10, wherein, wherein the plurality of operations further comprise maintaining the fan pitch below aerodynamic and operability limits.

12. The system of claim 8, wherein, wherein the first flight condition is a steady state cruise condition.

13. The system of claim 8, wherein, wherein the second flight condition is an increase power transient.

14. The system of claim 8, wherein, wherein the engine comprises at least one of a hybrid electric engine, an electric engine, or a fuel burning engine.

15. A method for controlling the clearance between a plurality of rotor blades and a shroud assembly of an aircraft engine, characterized by, The method comprises: receiving, with 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 total thrust constant of the aircraft engine, wherein adjusting the fan pitch includes implementing a first pitch response for a first portion of the step climb, and subsequently implementing a second pitch response for a second portion of the step climb, in which the fan pitch is increased to aerodynamic and operability limits.

16. The method of claim 15, wherein, wherein the power parameter includes at least one of a fan speed, an engine pressure ratio, or a cooling flow.

17. The method of claim 15, wherein, further comprising independently controlling the power parameter and the fan pitch.

18. The method of claim 15, wherein, further comprising maintaining the fan pitch below the aerodynamic and operability limits.

19. The method of claim 15, wherein, wherein the aircraft engine includes at least one of a hybrid-electric engine, an electric engine, or a fuel burning engine.

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