Gas turbine engine with clearance-based optimized acceleration rate

Abstract

A method of operating a gas turbine engine includes operating the gas turbine engine under a first flight condition. The method also includes receiving a demand for a second flight condition different from the first flight condition. The method further includes determining a final target gap and a transient target gap between the first component and the second component. The final target gap is associated with a second flight condition. The method also includes adjusting the engine acceleration rate to a nominal acceleration rate. The method also includes comparing the actual gap to the transient target gap after the time increment. The method also includes adjusting an engine acceleration rate based on a comparison between the actual gap and the transient target gap, wherein the transient target gap is adjusted based at least in part on the adjustment of the engine acceleration rate.

Description

Technical Field

[0001] The present subject matter relates generally to gas turbine engines. More specifically, the present subject matter relates to clearance control techniques for gas turbine engines. Background Art

[0002] Gas turbine engines typically include a clearance control system that is configured to adjust the clearance between the rotating turbine blades and the stationary casing to optimize efficiency and prevent contact between the turbine blades and the stationary casing. However, in certain rapid acceleration / deceleration situations, known clearance control systems cannot react quickly enough to prevent extrusion or friction events. For example, when a gas turbine engine needs to accelerate or decelerate rapidly, the mechanical expansion rate of the turbine blades exceeds the thermal expansion rate of the casing. In many cases, the clearance control system is not fast enough to account for the mechanical expansion of the turbine blades, which may cause the rotor blades to rub or squeeze the casing of the gas turbine engine. This may potentially cause damage to the gas turbine engine or reduce the efficiency of the engine.

[0003] Since the mechanical growth of the turbine blades can be faster than the thermal response of the surrounding turbine casing (and faster than the response of the clearance control system), it is necessary for engine designers to account for additional clearance to prevent rubbing or pinching conditions in these situations. However, by increasing the clearance between the end of the rotor blade and the engine casing, more air is able to escape through the rotor blade rather than passing through the blade, resulting in reduced gas turbine engine performance and increased fuel burn.

[0004] Therefore, improved systems and methods that allow for rapid acceleration and / or deceleration of a gas turbine engine while minimizing clearances for optimal performance are desired and would be welcomed in the art. BRIEF DESCRIPTION OF THE DRAWINGS

[0005] A full and enabling disclosure of the present 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:

[0006] Figure 1 A schematic cross-sectional view of a gas turbine engine according to an example embodiment of the present disclosure is provided;

[0007] Figure 2 A schematic cross-sectional view of another gas turbine engine according to an example embodiment of the present disclosure is provided;

[0008] Figure 3 Provided Figure 1 A close-up cross-sectional view of the aft end of the combustion section and the HP turbine of a gas turbine engine;

[0009] Figure 4is a schematic block diagram of an integrated engine control and flight operations system according to an exemplary embodiment of the present disclosure.

[0010] Figure 5 A control logic diagram according to an embodiment of the present disclosure is shown, which control logic may be implemented by an engine controller for controlling the clearance between turbine blades and a casing by adjusting the acceleration rate of a gas turbine engine;

[0011] Figure 6 Another control logic diagram according to an embodiment of the present disclosure is shown, which control logic may be implemented by an engine controller for controlling the clearance between turbine blades and a casing by adjusting a deceleration rate of a gas turbine engine;

[0012] Figure 7 is a graph of engine speed versus time according to an exemplary aspect of the present disclosure.

[0013] Figure 8 is a gap response graph of gap versus time, which shows how the gap in the gas turbine engine responds Figure 7 Adjustment of the acceleration shown.

[0014] Fig. 9 is a graph of engine speed versus time according to an exemplary aspect of the present disclosure.

[0015] Fig.10 is a gap response graph of gap versus time, which shows how the gap in the gas turbine engine responds Figure 7 Adjustment of the acceleration rate shown.

[0016] Fig.11 A flow chart of a method of operating a gas turbine engine according to an embodiment of the present disclosure is provided.

[0017] Fig.12 A block diagram of an engine controller according to an example embodiment of the present disclosure is provided. DETAILED DESCRIPTION

[0018] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numbers in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0019] 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. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0020] As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0021] The terms "upstream" and "downstream" refer to the relative flow directions with respect to the flow of a fluid in a fluid path. For example, "upstream" refers to the flow direction from which the fluid is flowing, and "downstream" refers to the flow direction toward which the fluid is flowing. "HP" means high pressure, and "LP" means low pressure.

[0022] Unless otherwise indicated herein, 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 via one or more intermediate components or features.

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

[0024] The term "at least one" in a context such as "at least one of A, B, and C" means only A, only B, only C, or any combination of A, B, and C.

[0025] Approximate language as used in the entire specification and claims herein is applied to modify any quantitative representation that can allow variation without causing the basic function associated therewith to change. Therefore, the value modified by 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 applied to a single value, any endpoint or two endpoints that limit the range of values, and / or the margin of the range between the endpoints.

[0026] The present disclosure generally relates to preventing extrusion between rotor blades and a casing of a gas turbine engine of an aircraft. Under certain operating conditions, such as step climb conditions or other conditions, the gas turbine engine may need to accelerate or decelerate rapidly, which results in mechanical and thermal expansion of the rotor blades and thermal expansion of the casing surrounding the rotor blades. Under such rapid acceleration / deceleration conditions, the mechanical expansion rate of the turbine blades exceeds the thermal expansion rate of the casing, which may cause the rotor blades to rub or squeeze the casing of the gas turbine engine. This may potentially result in damage to the gas turbine engine or reduce the efficiency of the engine.

[0027] Since the mechanical growth of the rotor blades can be faster than the thermal reaction of the surrounding turbine casing, it is necessary for engine designers to consider additional clearance to prevent rubbing or pinching conditions in these situations. However, by increasing the clearance between the end of the rotor blade and the engine casing, more air is able to escape through the rotor blade rather than passing through the blade, resulting in reduced gas turbine engine performance and increased fuel burn.

[0028] Specifically, as used herein, "acceleration rate" refers to the angular acceleration value at a certain moment (e.g., the time rate of change of the angular velocity of a specific rotating component (such as the LP axis, HP axis, etc.)). That is, "acceleration rate" does not refer to the rate of change of acceleration (e.g., jerk), but refers to the acceleration value at that moment. For the purpose of this disclosure, "acceleration rate" and "acceleration" can be used interchangeably.

[0029] The present disclosure relates to systems and methods for continuously linking the acceleration rate of a gas turbine engine to a target clearance. For example, when the gas turbine engine is operating at a first flight condition (such as a steady-state condition) and a demand is received to operate at a second flight condition (such as a step climb condition), the gas turbine engine can be set to the final target clearance to be achieved at the second flight condition and can begin to accelerate or decelerate. At one or more points during the acceleration or deceleration to the second flight condition, the system can determine whether the gas turbine engine is on pace to achieve the final target clearance associated with the second flight condition. If not, the system can adjust the acceleration or deceleration of the gas turbine engine. If so, the system can continue to accelerate or decelerate at the current rate without any adjustments.

[0030] Now referring to the accompanying drawings, Figure 1 A schematic cross-sectional view of a gas turbine engine 100 is provided in accordance with an example embodiment of the present disclosure. Figure 1 In the depicted embodiment, the gas turbine engine 100 is an aviation high bypass turbofan jet engine that is configured to be mounted to an aircraft, for example, in an underwing configuration. As shown, the gas turbine engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. The axial direction A extends parallel to or coaxially with a longitudinal centerline 102 defined by the gas turbine engine 100.

[0031] The gas turbine engine 100 includes a fan section 104 and a core turbine engine 106 disposed downstream of the fan section 104. The core turbine engine 106 includes an engine cowling 108 defining an annular inlet 110. The engine cowling 108 encloses, in serial flow relationship: a compressor section 112 including a first boost or LP compressor 114 and a second HP compressor 116; a combustion section 118; a turbine section 120 including a first HP turbine 122 and a second LP turbine 124; and an exhaust section 126. An HP shaft 128 drivingly connects the HP turbine 122 to the HP compressor 116. An LP shaft 130 drivingly connects the LP turbine 124 to the LP compressor 114. The compressor section 112, the combustion section 118, the turbine section 120, and the exhaust section 126 together define a core air flow path 132 through the core turbine engine 106.

[0032] The fan section 104 includes a fan 134 having a plurality of fan blades 136 coupled to a disk 138 in a circumferentially spaced manner. As depicted, the fan blades 136 extend outwardly from the disk 138 generally in a radial direction R. Each fan blade 136 can rotate relative to the disk 138 about a pitch axis P by means of the fan blades 136 being operably coupled to a suitable actuating member 140, the actuating member 140 being configured to jointly change the pitch of the fan blades 136, for example, in unison. The fan blades 136, the disk 138, and the actuating member 140 can rotate together about the longitudinal centerline 102 via the LP shaft 130 across a power gearbox 142. The power gearbox 142 includes a plurality of gears for reducing the rotational speed of the LP shaft 130 to affect a more efficient rotating fan speed. In other embodiments, the fan blades 136, the disk 138, and the actuating member 140 can be directly connected to the LP shaft 130, for example, in a direct drive configuration. Furthermore, in other embodiments, fan blades 136 of fan 134 may be fixed pitch fan blades.

[0033] Still reference Figure 1 , disk 138 is covered by a rotatable spinner 144 that is aerodynamically shaped to facilitate airflow through a plurality of fan blades 136. In addition, fan section 104 includes an annular fan case or outer nacelle 146 that circumferentially surrounds fan 134 and / or at least a portion of core turbine engine 106. Nacelle 146 is supported relative to core turbine engine 106 by a plurality of circumferentially spaced outlet guide vanes 148. A downstream section 150 of nacelle 146 extends over an outer portion of core turbine engine 106 to define a bypass airflow passage 152 therebetween.

[0034] During operation of gas turbine engine 100, a volume of air 154 enters gas turbine engine 100 through nacelle 146 and / or an associated inlet 156 of fan section 104. As volume of air 154 passes through fan blades 136, a first portion of air 154 (as indicated by arrow 158) is directed or directed into bypass airflow passage 152, and a second portion of air 154 (as indicated by arrow 160) is directed or directed into LP compressor 114. As second portion of air 160 is directed through LP compressor 114 and HP compressor 116, the pressure of second portion of air 160 increases. Compressed second portion of air 160 is then discharged into combustion section 118.

[0035] The compressed second portion of air 160 from the compressor section 112 is mixed with fuel and combusted within the combustor of the combustion section 118 to provide combustion gases 162. For example, the fuel may be supplied from a fuel supply system 200. The fuel supply system 200 may include a fuel storage device 201, a fuel supply line 202, and a control valve 204, which is disposed in fluid communication on the fuel supply line 202. The control valve 204 is operable to adjust the amount of fuel supplied to the combustion section 118. The combustion gases 162 are directed from the combustion section 118 along the hot gas path 174 of the core air flow path 132 through the HP turbine 122, wherein a portion of the thermal energy and / or kinetic energy from the combustion gases 162 is extracted via sequential stages of HP turbine stator vanes 164 and HP turbine blades 166. The HP turbine blades 166 are mechanically coupled to the HP shaft 128. Thus, when the HP turbine blades 166 extract energy from the combustion gases 162, the HP shaft 128 rotates, which supports the operation of the HP compressor 116. The combustion gases 162 are directed through the LP turbine 124, wherein a second portion of thermal and kinetic energy is extracted from the combustion gases 162 via sequential stages of LP turbine stator vanes 168 and LP turbine blades 170. LP turbine blades 170 are coupled to the LP shaft 130. Thus, when LP turbine blades 170 extract energy from the combustion gases 162, the LP shaft 130 rotates, which supports the operation of the LP compressor 114 and the fan 134.

[0036] The combustion gases 162 are then directed through the exhaust section 126 of the core turbine engine 106 to provide propulsive thrust. At the same time, the pressure of the first portion of air 158 is significantly increased, also providing propulsive thrust, as it is directed through the bypass airflow passage 152 before being discharged from the fan nozzle exhaust section 172 of the gas turbine engine 100. The HP turbine 122, the LP turbine 124, and the exhaust section 126 at least partially define a hot gas path 174 for directing the combustion gases 162 through the core turbine engine 106.

[0037] like Figure 1 As further shown, the gas turbine engine 100 includes a clearance adjustment system, which in this embodiment is an active clearance control (ACC) system 101. Generally, the ACC system 101 is configured to dynamically control blade tip clearance between a rotating component (e.g., a turbine blade) and a stationary component (e.g., a shroud). For this embodiment, the ACC system 101 includes one or more compressor supply ducts, such as a compressor supply duct 195 that feeds into a supply duct 191. As shown, the supply duct 191 provides a conduit for hot control air 197 to flow from the HP compressor 116 of the compressor section 112 to the HP turbine 122 and / or the LP turbine 124. Additionally or alternatively, although in Figure 1 Not shown in the example embodiment of FIG. 1 , supply duct 191 may be configured to deliver air from fan section 104 and / or LP compressor 114 to HP turbine 122 and / or LP turbine 124 .

[0038] The mass flow and temperature of hot control air 197 provided to the HP turbine 122 and / or the LP turbine 124 are controlled by adjusting the first control valve 192 and / or the second control valve 193. For this embodiment, the first control valve 192 controls the bleed air from the HP compressor 116 to the HP turbine 122 when adjusted. In other embodiments, the source air for the ACC system 101 may come from in front of the compressor section 112, immediately behind the fan section 104, or other locations. The second control valve 193 controls the bleed air from the HP compressor 116 to the LP turbine 124 when adjusted. The first control valve 192 and the second control valve 193 or controllable devices are controlled by one or more engine controllers and are in communication therewith. Figure 1 In the depicted embodiment, an engine controller 402 is housed within the nacelle 146. The controller 402 may be, for example, an electronic engine controller (EEC) or an electronic control unit (ECU) of a full authority digital engine control (FADEC) system. The engine controller 402 includes various components for performing various operations and functions, such as controlling clearances and / or accelerations of the gas turbine engine 100. The controller 402 (or computing system) may have one or more processing devices and one or more memory devices. The one or more memory devices, such as one or more non-transitory computer-readable media, may store computer-readable instructions that may be executed by the one or more processing devices to perform operations, such as controlling operation of the gas turbine engine 100. The controller 402 may be configured as follows: Fig.12 shown.

[0039] In many embodiments, the engine controller 402 is operable to control the acceleration (i.e., the rate of change of the rotational speed of the HP shaft 128 and / or the LP shaft 130) of the gas turbine engine 100. The engine controller may actuate the control valve 204 to adjust the amount of fuel supplied to the combustion section 118 to increase or decrease the acceleration of the gas turbine engine 100. For example, when the control valve 204 is open, more fuel is supplied to the combustion section 118, which increases the acceleration of the HP shaft 128 and / or the LP shaft 130 and the entire gas turbine engine 100.

[0040] When the control valves 192, 193 are open, relatively cool or hot hot control air 197 flows from the HP compressor 116 to the HP turbine 122 and the LP turbine 124. When the hot control air 197 reaches the HP turbine 122, a distribution manifold 175 associated with the HP turbine 122 distributes the hot control air 197 around the HP turbine 122 so that blade tip clearance can be controlled. When the hot control air 197 reaches the LP turbine 124, a distribution manifold 177 associated with the LP turbine 124 distributes the hot control air 197 around the LP turbine 124 so that blade tip clearance can be controlled. When the control valves 192, 193 are closed, the hot control air 197 is restricted from flowing to the HP turbine 122 and the LP turbine 124, which results in minimal control air being supplied to the housing. When one of the control valves 192, 193 is open and the other is closed, the hot control air 197 is allowed to flow to the turbine associated with the open control valve while restricting the hot control air 197 from flowing to the turbine associated with the closed control valve.

[0041] although Figure 1 The embodiment of is shown as having two control valves 192, 193, but it should be understood that any suitable number of control valves may be included. In some alternative embodiments, such as Figure 2 As depicted, the ACC system 101 may include a single control valve 194 that selectively allows hot control air 197 to flow to the HP turbine 122 and the LP turbine 124. In such an embodiment, as shown, the control valve 194 may be disposed on (and in fluid communication with) a compressor supply conduit 195 (e.g., upstream of the supply conduit 191). In other embodiments, one or more control valves may be located along a supply conduit configured to deliver air from the fan section 104 to the HP turbine 122 and / or the LP turbine 124. Other configurations are also possible.

[0042] Furthermore, it should be understood that Figure 1The ACC system 101 depicted in FIG. 1 is an example gap adjustment system. In other example embodiments, the gap adjustment system may have other suitable configurations. For example, in some embodiments, the gap adjustment system may include one or more electric heating elements, without cooling air or with fixed cooling air to adjust the gap. Other gap adjustment systems are contemplated.

[0043] Furthermore, it should be understood that Figure 1 The gas turbine engine 100 depicted in the drawings is provided as an example only, and in other example embodiments, the gas turbine engine 100 may have any other suitable configuration. Additionally or alternatively, aspects of the present disclosure may be used with any other suitable aviation gas turbine engine (e.g., a turboshaft engine, a turboprop engine, a turbojet engine, etc.). Furthermore, aspects of the present disclosure may also be used with any other land-based gas turbine engine (e.g., a power generation gas turbine engine) or any aero-derivative gas turbine engine (e.g., a marine gas turbine engine).

[0044] Figure 3 Provided Figure 1 A close-up cross-sectional view of the aft end of the combustion section 118 and the HP turbine 122 of the gas turbine engine 100 is shown. Figure 3 , the HP turbine 122 includes a first stage 176 in a serial flow relationship, the first stage 176 including an annular array 178 of stator vanes 164a (only one shown) axially spaced from an annular array 180 of turbine blades 166a (only one shown). The HP turbine 122 also includes a second stage 182, the second stage 182 including an annular array 184 of stator vanes 164b (only one shown) axially spaced from an annular array 186 of turbine blades 166b (only one shown). The turbine blades 166a, 166b extend radially from the HP shaft 128 and are coupled to the HP shaft 128 by rotor disks 167a, 167b. The stator vanes 164a, 164b and the turbine blades 166a, 166b direct the combustion gases 162 from the combustion section 118 through the HP turbine 122 along a hot gas path 174.

[0045] like Figure 3As further depicted in , the HP turbine 122 includes shroud assemblies 188a, 188b, each shroud assembly forming an annular ring around an annular array of blades. In particular, the shroud assembly 188a forms an annular ring around the annular array 180 of blades 166a of the first stage 176, and the shroud assembly 188b forms an annular ring around the annular array 186 of turbine blades 166b of the second stage 182. For this embodiment, the shroud assemblies 188a, 188b include shrouds 190a, 190b coupled to respective hangers 196a, 196b, which in turn are coupled to a turbine casing 198.

[0046] The shrouds 190a, 190b of the shroud assemblies 188a, 188b are radially spaced from the blade tips 192a, 192b of the turbine blades 166a, 166b. Blade tip clearances CL are defined between the blade tips 192a, 192b and the shrouds 190a, 190b. It should be noted that blade tip clearances CL may similarly exist in the LP compressor 114, the HP compressor 116, and / or the LP turbine 124. Thus, the subject matter disclosed herein is not limited to adjusting blade tip clearances and / or clearance closures in HP turbines; rather, the teachings of the present disclosure may be used to adjust blade tip clearances in any suitable section of the gas turbine engine 100.

[0047] As previously described, the ACC system 101 regulates a relatively cool or hot flow of thermal control air 197 from the fan section 104 and / or the compressor section 112 and disperses the air over the HP and / or LP turbine casings (e.g., the turbine casing 198 of the HP turbine 122) to contract or expand the turbine casing relative to the HP / LP turbine blade tips depending on, among other factors, aircraft and engine operating and flight conditions. Figure 3 As shown, hot control air 197 is directed to the HP turbine 122 via supply conduit 191. In some embodiments, the hot control air 197 can be directed through a heat exchanger (not shown) to further cool or heat the air. The hot control air 197 enters the distribution manifold 175 through an inlet 199 defined by the distribution manifold 175. The hot control air 197 is distributed on the turbine casing 198 via the distribution manifold 175. In this way, the blade tip clearance CL can be controlled. The amount of hot control air 197 provided to the HP turbine 122 (and / or the LP turbine 124) can be controlled by adjusting the control valves 192, 193 ( Figure 1 ) to control.

[0048] It will be appreciated that engine performance depends at least in part on the blade tip clearance CL between the turbine blade tip and the shroud. Generally, the tighter the clearance between the blade tip and the shroud (i.e., the more closed the clearance), the more efficient the gas turbine engine operates. Therefore, minimizing or otherwise reducing the blade tip clearance CL helps to optimize and / or otherwise improve engine performance and efficiency. However, one challenge in minimizing the blade tip clearance CL is that turbine blades expand and contract at a different rate than the shroud and casing that circumferentially surround them.

[0049] More specifically, the blade tip clearance CL between the tip of the turbine blade and the surrounding shroud and turbine casing may be affected by two main types of loads: power-induced engine loads and flight loads. Power-induced engine loads typically include centrifugal loads, thermal loads, internal pressure loads, and thrust loads. Flight loads typically include inertial loads, aerodynamic loads, and gyroscopic loads. Centrifugal loads and thermal engine loads are the cause of the largest radial variation in blade tip clearance CL. With respect to centrifugal loads, the blades of a turbine engine may mechanically expand or contract depending on their rotational speed. Generally, the faster the rotational speed of the rotor, the greater the mechanical expansion of the turbine blades, and therefore the further the blades extend radially outward. Conversely, the slower the rotational speed of the rotor, the smaller the mechanical expansion experienced by the rotor, and the further the blades extend radially inward from the centerline longitudinal axis of the engine. With respect to thermal loads, when the engine heats up or cools down at least in part due to changes in power levels (i.e., changes in engine speed), the rotor and casing thermally expand and / or contract at different rates. That is, the rotor is relatively large and heavy, so the thermal mass of the rotor heats up and cools down at a much slower rate than the relatively thin and light turbine casing. Therefore, the thermal mass of the housing heats up and cools down faster than the rotor.

[0050] Therefore, when the aircraft maneuvers and its engine undergoes various power level changes, the rotor and the housing contract and expand at different rates. Therefore, the rotor and the housing are sometimes not thermally matched. This mismatch can cause changes in the blade tip clearance CL, and in some cases, the turbomachinery components may contact or rub against each other, resulting in a friction event. For example, a friction event may occur where the blade tips 192a, 192b contact or touch the corresponding shrouds 190a, 190b. Friction events may cause poor engine performance and efficiency, may shorten the effective service life of the turbine blades 166a, 166b and / or shrouds 190a, 190b, and may degrade the exhaust temperature margin of the engine. Therefore, ideally, the blade tip clearance CL is set to minimize the gap between the blade tip and the shroud without subjecting the turbomachinery components to friction events. Taking these aspects into account, this article provides a control technology for setting the clearance.

[0051] Figure 4is a schematic block diagram of an integrated engine control and flight operation system 400 according to an exemplary embodiment of the present disclosure. In the exemplary embodiment, the integrated system 400 includes an engine controller 402, such as but not limited to a FADEC, mounted proximate to an associated aircraft engine or gas turbine engine 100. The gas turbine engine 100 includes a fan 410 and a core engine 412 in serial flow communication. In some embodiments, substantially all of the air flow through the fan 410 passes through the core engine 412. In various embodiments, the gas turbine engine 100 is a high bypass type engine, and only a portion of the air flow entering the fan 410 passes through the core engine 412. Although described as a FADEC, in various embodiments, the engine controller 402 may include other forms of engine controllers capable of operating as described herein.

[0052] A plurality of process sensors 414 are positioned about the gas turbine engine 100 to sense process parameters associated with the gas turbine engine 100. Such process parameters include, for example, engine speed, fuel flow, damper and guide vane positions, stator vane clearance, and various temperatures of components in the gas turbine engine 100. The sensors 414 are communicatively coupled to the engine controller 402. For example, captured values ​​or sensor data 415 may be directed to the engine controller 402. One or more sensors 414 may capture operating parameter values ​​continuously, may capture operating parameter values ​​at predetermined intervals, and / or when conditions are met.

[0053] In some embodiments, the one or more sensors 414 may include at least one sensor operable to directly measure a clearance between a rotating component and a stationary component of the gas turbine engine 100. For example, the one or more sensors 414 may include a sensor 232a operable to measure a clearance between a turbine blade 166a and a shroud 190a. Figure 3 The one or more sensors 414 may also include a sensor 232b operable to measure a clearance between the turbine blades 166b and the shroud 190b ( Figure 3 ). The sensors 232a, 232b may be optical probes, inductive proximity sensors, combinations thereof, or any suitable type of sensor operable to directly measure the clearance between their respective rotating and stationary components. The sensors 232a, 232b may each capture the instantaneous clearance between their respective turbine blades 166a, 166b and shrouds 190a, 190b and may provide the instantaneous clearance or measured clearance as part of the sensor data 415 to the engine controller 402.

[0054] The one or more sensors 414 may also include at least one sensor operable to directly measure a clearance between a rotating component and a stationary component of the LP turbine 124. The sensor positioned in the LP turbine 124 may capture an instantaneous clearance between the LP turbine blades 170 (or an array of LP turbine blades) and their associated shrouds and may provide the instantaneous clearance or measured clearance to the engine controller 402 as part of the sensor data 415.

[0055] The one or more sensors 414 may also include other sensors. The one or more sensors 414 may include sensors operable to capture or measure operating parameter values ​​of various operating parameters (e.g., various speeds, pressures, temperatures, etc.) indicative of an operating condition or operating point of the gas turbine engine 100. Example operating parameters include, but are not limited to, shaft speed of the LP shaft 130, shaft speed of the HP shaft 128, compressor discharge pressure, ambient temperature, ambient pressure, temperature along the hot gas path 174 between the HP turbine 122 and the LP turbine 124, the altitude at which the gas turbine engine 100 operates, etc. Such sensors may measure or capture operating parameter values ​​of their respective operating parameters, and such operating parameter values ​​may be directed to the engine controller 402 as part of the sensor data 415. The sensor data 415 may also include data indicating a power level of the gas turbine engine 100, for example, based on a position of a throttle of the gas turbine engine 100.

[0056] In many embodiments, the sensor 414 is operable to capture or measure the acceleration of the gas turbine engine 100. That is, the rate of change of the speed at which the LP shaft 130 and / or the HP shaft 128 rotate. In other embodiments, one or more sensors 414 are operable to capture or measure the speed of the LP shaft 130 and / or the shaft speed of the HP shaft 128, and using the measured speeds, the engine controller 402 can determine the engine acceleration. In an exemplary embodiment, the engine controller 402 is operable to adjust the engine acceleration, such as by adjusting (i.e., increasing or decreasing) the amount of fuel supplied to the combustion section 118.

[0057] In addition, one or more actuators 416 are positioned around the gas turbine engine 100 and are operably coupled to components of the gas turbine engine 100 to enable operation of these components. The actuators 416 may also be communicatively coupled to the engine controller 402. In one embodiment, the actuators 416 are in communication with the control valves 192, 193, 194, 204. The actuators 416 may receive commands to open or close one or more control valves 192, 193, 194 to control the heating / cooling flow delivered according to the engine response plan, as further described herein. In addition, the actuators 416 may receive commands to adjust the amount of fuel supplied to the combustion section 118 by actuating the control valve 204, which adjusts the acceleration of the gas turbine engine. The engine controller 402 uses the sensors 414 and the actuators 416 to determine the operating conditions of the gas turbine engine 100, including but not limited to the performance of the gas turbine engine 100 relative to a baseline or new operating conditions. Engine controller 402 may then operate actuators 416 to account for degradation and / or damage to gas turbine engine 100 between overhauls. Engine controller 402 may also use sensors 414 and actuators 416 to store determined engine conditions for future reference, further processing, and / or reporting.

[0058] The system 400 also includes a flight control system 420 (eg, a flight management system or FMS) communicatively coupled to the engine controller 402 via a communication channel 422. The flight control system 420 includes a processor 421 and a memory 423 communicatively coupled to the processor 421. Figure 4 In an exemplary embodiment of the present invention, the communication channel 422 is a wired connection between the engine controller 402 and the flight control system 420. In various other embodiments, the communication channel 422 may be a wireless communication medium. In an exemplary embodiment, the flight control system 420 is positioned close to the cockpit (not shown) of the aircraft, and the engine controller 402 is positioned close to the engine associated therewith. The flight control system 420 may be embodied in a single processor-based component, or the functions of the flight control system 420 may be performed by multiple components configured to perform the functions described herein. For convenience, safety, and / or optimal operation considerations, some of the components that perform the functions of the flight control system 420 may be positioned close to the cockpit, while other components may be distributed inside the aircraft. 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 the engine controller and any avionics functions installed on the aircraft.

[0059] The flight control system 420 is configured to interface with various other systems on and off the aircraft. For example, the flight control system 420 may receive current aircraft status from a plurality of aircraft sensors 424 via a sensing system 426. Such sensors 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, state, or position of the aircraft, such as aircraft orientation, time to target, or other conditions. The flight control system 420 may also receive information from one or more onboard processing systems 428, which may be stand-alone systems or systems having functionality distributed across several computer systems. The flight control system 420 and the onboard processing system 428 may use wired communication channels and / or network connections (e.g., Ethernet or fiber optics), wireless communication devices (such as radio frequency (RF) (e.g., FM radio and / or digital audio broadcasting)), Institute of Electrical and Electronics Engineers (IEEE) communications. 802.11 standards (e.g., 802.11(g) or 802.11(n)), Worldwide Interoperability for Microwave Access The flight control system 420 may communicate using one or more wireless communication media, including, but not limited to, radio frequency (RF) (e.g., FM radio and / or digital audio broadcasting), Institute of Electrical and Electronics Engineers (IEEE), wireless communications technology, and / or other suitable communication means. As used herein, a wired communication channel includes channels that communicate using optical fiber and other optical means. The flight control system 420 may also receive information from one or more off-board processing systems 430, which may be stand-alone systems or systems having functionality distributed across several computer systems and / or several sites. The off-board processing systems 430 and the flight control system 420 may communicate using one or more wireless communication media, including, but not limited to, radio frequency (RF) (e.g., FM radio and / or digital audio broadcasting), Institute of Electrical and Electronics Engineers (IEEE), wireless communications technology, and / or other suitable communication means. 802.11 standards (such as 802.11(g) or 802.11(n)), Worldwide Interoperability for Microwave Access standards, cellular telephone technology (eg, Global Standard for Mobile Communications (GSM)), satellite communication links), and / or any other suitable communication means.

[0060] As in at least some aircraft operating procedures, a step climb condition occurs when the pilot of the aircraft chooses to increase the altitude at which the aircraft is traveling. According to current FAA regulations, altitude steps usually occur in 2,000-foot increments. This means, for example, that a pilot of an aircraft flying at 33,000 feet can choose to perform a step climb maneuver to cause the aircraft to climb 2,000 feet to an altitude of 35,000 feet. In order to achieve a step climb, the pilot modifies the control of the autopilot / autothrottle system of the flight control system 420 to request that the aircraft rise to a desired cruising altitude. The flight control system 420 then uses a predetermined algorithm to increase engine power in order to cause the aircraft to climb. Because the request for increased engine power usually requires the gas turbine engine 100 to accelerate, which increases the engine thrust, the turbine blades 166 (e.g., rotor blades) grow due to mechanical forces and associated thermal changes. This turbine blade growth results in a reduction in clearance. If the growth exceeds the design tolerance, the turbine blades may rub or squeeze the casing of the gas turbine engine 100, potentially causing damage or reducing the efficiency of the engine.

[0061] For the ACC system 101, the thermal growth of the gas turbine engine 100 casing can be matched to the thermal and mechanical growth of the turbine blades 166 if sufficient time is given for the clearance control system to operate. For example, a step climb from approximately 33,000 feet to 35,000 feet may require the aircraft to take more than two minutes to complete. However, known flight control system step climb algorithms command the gas turbine engine 100 to respond to a request for increased thrust within, for example, 5 seconds, resulting in the growth rate of the turbine blades 166 exceeding the growth rate of the turbine casing 198. Because the mechanical growth of the turbine blades 166 is faster than the thermal reaction of the surrounding turbine casing 198, it is necessary for the designer of the engine 100 to consider additional clearances to prevent rubbing or crushing conditions in these situations (e.g., in situations where the mechanical growth of the turbine blades 166 is faster than the acceleration or deceleration of the casing 198). In the above example, the additional clearance is referred to herein as step climb clearance. However, by increasing the clearance between the ends of turbine blades 166 and engine casing 198 , more air is able to escape through the turbine blades rather than traveling through the blades, resulting in reduced performance of gas turbine engine 100 and increased fuel burn.

[0062] Figure 5 and Figure 6 1 and 2 respectively illustrate control logic 500 and 600 that may be implemented by the engine controller 402 for controlling the relationship between the blade tips 192a, 192b and the shrouds 190a, 190b ( Figure 3) between the blade tip clearance (such as blade tip clearance CL). The engine controller 402 may implement the control logic 500 and / or 600 in response to a request for a new flight condition. For example, when the gas turbine engine 100 is operating in a first flight condition (such as a steady-state condition), and the engine controller 402 receives a request for a second flight condition (such as a step climb condition, an acceleration condition, or a deceleration condition), the engine controller 402 may implement the control logic 500 or 600.

[0063] Specific reference Figure 5 In response to receiving a request for a new flight condition, the control logic may include, at (502), an initial step of beginning to accelerate the gas turbine engine at a nominal acceleration rate. The nominal acceleration rate may be a predetermined acceleration rate that the engine controller 402 may implement by default until a condition for an optimized acceleration rate is marked based on a transient target clearance. The nominal acceleration rate may include an HP nominal acceleration rate for the HP shaft 128 and an LP nominal acceleration rate for the LP shaft 130. The HP nominal acceleration rate may be the same as or different from the LP nominal acceleration rate.

[0064] In many embodiments, the control logic may include determining at (504) whether a gas turbine engine 100 operating condition for which an optimized acceleration rate (e.g., a deviation or modification from a nominal acceleration rate) is marked. Many operating conditions of the gas turbine engine 100 for which an optimized acceleration rate may be marked may include a maximum acceleration condition (when the gas turbine engine accelerates from a first speed to a second speed), a step climb condition, or others. However, under some operating conditions, it may be optimal to proceed at the nominal acceleration rate. For example, if the aircraft is very close to a runway (such as a takeoff or landing condition). The engine controller 402 may identify when the operating conditions of the gas turbine engine 100 are optimal for applying the optimized acceleration rate to ensure that the optimized acceleration rate is not applied at an inappropriate time.

[0065] When the engine controller 402 determines that the operating conditions of the gas turbine engine 100 are such that the optimized acceleration rate can be implemented, the control logic diagram 500 can include determining, using the engine controller 402 at (506), whether the calculated, measured, and / or sensed clearance is consistent with the transient target clearance. In other words, the engine controller 402 can align the blade tips 192a, 192b with the shrouds 190a, 190b ( Figure 3 ) is compared with a target gap at a specific point in time.

[0066] When the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3) is within a predetermined margin of the target gap at that particular point in time (e.g., within a ±15% margin, or such as within a ±10% margin, or such as within a ±5% margin, or such as within a ±1% margin), the control logic may include continuing at the current acceleration rate at (508) (i.e., without modifying or adjusting the acceleration of the gas turbine engine 100).

[0067] However, when the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3 ) is outside a predetermined margin of the target clearance at that particular point in time (e.g., outside a ±15% margin, or such as outside a ±10% margin, or such as outside a ±5% margin, or such as outside a ±1% margin), then the control logic may include adjusting the acceleration rate at (510) to achieve the transient target clearance. For example, when the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3 ) is outside a predetermined margin of the target clearance at that particular point in time, the acceleration rate can be increased or decreased to achieve the transient target clearance. In one example, when it is determined that the actual clearance exceeds a predetermined margin of the transient target clearance (e.g., greater than an upper threshold of the predetermined margin, i.e., the clearance is large and may result in low efficiency), the engine acceleration rate can be increased until the actual clearance is within a predetermined margin of the transient clearance. In another example, when it is determined that the actual clearance is less than a predetermined margin of the transient target clearance (e.g., less than a lower threshold of the predetermined margin, i.e., the clearance is too small and may result in squeezing / rubbing), the engine acceleration rate can be decreased until the actual clearance is within a predetermined margin of the transient clearance. In other words, when the gas turbine engine is in a transient state (i.e., when the gas turbine engine 100 transitions from a first operating condition to a second operating condition by accelerating / decelerating), the engine controller 402 can periodically (or in some embodiments, continuously) determine the relative position of the blade tips 192a, 192b with the shrouds 190a, 190b ( Figure 3 ) is within a predetermined margin (such as within ±5%, or within ±10%, or within ±20%) of the target gap at that particular moment (if not, the engine controller can adjust the acceleration rate).

[0068] In various cases, such as Figure 5As shown, before repeating decision steps (504) and / or (506), the engine controller 402 may wait for a time increment at (505). The time increment may be a standard increment, such as between about 60 seconds and about 0.1 seconds, or such as between about 30 seconds and about 0.1 seconds, or such as between about 10 seconds and about 0.1 seconds, or such as between about 1 second and about 0.1 seconds. In some embodiments, the time increment may be zero, so that the logic loops continuously until the target clearance is achieved. In some embodiments, the time increment may be a recalculation rate of the engine controller, which may be between about 0.05 seconds and about 0.4 seconds. In many embodiments, the control logic may iterate steps (506), (508) or (510) and (505) until the gas turbine engine 100 is operating at the second flight condition and / or the gas turbine engine has achieved the final target clearance. Additionally, the engine controller may iterate steps (502), (504), and (505) until the gas turbine engine 100 is operating at the second flight condition and / or the gas turbine engine 100 achieves the final target clearance. In some embodiments, the iterations may be performed in set steps, which may be equal to the time increment. In other embodiments, the iterations may be continuous.

[0069] Specific reference Figure 6 In response to receiving a request for a new flight condition, the control logic may include, at (602), an initial step of beginning to decelerate the gas turbine engine at a nominal deceleration rate. The nominal deceleration rate may be a predetermined deceleration rate that the engine controller 402 may implement by default until a condition for an optimized deceleration rate is flagged. The nominal deceleration rates may include an HP nominal deceleration rate for the HP shaft 128 and an LP nominal deceleration rate for the LP shaft 130. The HP nominal deceleration rate may be the same as or different from the LP nominal deceleration rate.

[0070] In many embodiments, the control logic may include determining at (604) whether a gas turbine engine 100 operating condition for an optimized deceleration rate (e.g., a deviation or modification from a nominal deceleration rate) is marked in order to maintain proper clearance. Many operating conditions of the gas turbine engine 100 for which an optimized deceleration rate may be marked, such as a maximum deceleration condition (when the gas turbine engine is decelerating from a first speed to a second speed). In addition, when a hot rotor and reacceleration are required, the engine controller 402 may mark an operating condition for an optimized deceleration rate in order to reduce the likelihood of a crush or friction event during reacceleration of the gas turbine engine 100. For example, as described above, the rotor and the casing thermally expand and / or contract at different rates. That is, the rotor is relatively large and heavy, and therefore, the thermal mass of the rotor heats and cools at a much slower rate than the relatively thin and light turbine casing. Therefore, the thermal mass of the casing heats and cools faster than the rotor. A hot rotor reacceleration situation may occur after deceleration of the gas turbine engine. For example, during deceleration, the casing cools and thermally contracts while the rotor remains hot and thermally expands, such that when the gas turbine engine re-accelerates, the turbine blades have less room for mechanical expansion, which may result in a pinch or rubbing event. The engine controller 402 may flag an optimized deceleration condition at (604) when a hot rotor re-acceleration condition is anticipated (i.e., the engine controller 402 determines that the gas turbine engine 100 will re-accelerate within some predetermined time period after the gas turbine engine 100 is decelerated).

[0071] However, under some operating conditions, it may be optimal to proceed at the nominal deceleration rate. For example, if the aircraft is very close to a runway (such as a takeoff or landing condition). The engine controller 402 can identify when the operating conditions of the gas turbine engine 100 are optimal for applying the optimized deceleration rate to ensure that the optimized deceleration rate is not applied at an inappropriate time.

[0072] When the engine controller 402 determines that the operating conditions of the gas turbine engine 100 are such that an optimized deceleration rate can be implemented, the control logic 600 may include determining, using the engine controller 402 at (606), whether the calculated, measured, and / or sensed clearance is consistent with the transient target clearance. In other words, the engine controller 402 may align the blade tips 192a, 192b with the shrouds 190a, 190b ( Figure 3 ) is compared with a target gap at a specific point in time.

[0073] When the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3) is within a predetermined margin of the target gap at that particular point in time (e.g., within a ±15% margin, or such as within a ±10% margin, or such as within a ±5% margin, or such as within a ±1% margin), the control logic may include continuing at the current deceleration rate at (608) (i.e., without modifying or adjusting the deceleration of the gas turbine engine 100).

[0074] However, when the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3 ) is outside a predetermined margin of the target clearance at that particular point in time (e.g., outside a ±15% margin, or such as outside a ±10% margin, or such as outside a ±5% margin, or such as outside a ±1% margin), then the control logic may include adjusting the deceleration rate at (610) to achieve the transient target clearance. In particular, in one non-limiting example, when the blade tips 192a, 192b are aligned with the shrouds 190a, 190b ( Figure 3 ) is greater than a predetermined margin of a target gap at a certain point in time, the deceleration rate can be increased.

[0075] In other words, when the gas turbine engine is in a transient state (i.e., when the gas turbine engine 100 transitions from a first operating condition to a second operating condition through acceleration / deceleration), the engine controller 402 may periodically (or in some embodiments, continuously) determine the position of the blade tips 192a, 192b relative to the shrouds 190a, 190b ( Figure 3 ) is within a predetermined margin of the target gap at that particular point in time (if not, the engine controller can adjust the deceleration rate).

[0076] In various cases, such as Figure 6As shown, before making decision steps (604) and / or (606), the engine controller 402 may wait for a time increment at (605). The time increment may be a standard increment, such as between about 60 seconds and about 0.1 seconds, or such as between about 30 seconds and about 0.1 seconds, or such as between about 15 seconds and about 0.1 seconds. In some embodiments, the time increment may be equal to the recalculation rate of the engine controller, which may be between about 0.05 seconds and about 0.4 seconds, or such as between about 0.1 seconds and about 0.3 seconds. In many embodiments, the control logic may iterate steps (606), (608) or (610) and (605) until the gas turbine engine 100 is operated at the second flight condition and / or the gas turbine engine has achieved the final target clearance. In addition, the engine controller may iterate steps (602), (604) and (605) until the gas turbine engine 100 is operated at the second flight condition and / or the gas turbine engine has achieved the final target clearance. In some embodiments, the iteration may be performed at a set step size, which may be equal to the time increment. In other embodiments, the iteration may be continuous.

[0077] Figure 7 and Figure 8 The above reference is graphically depicted Figure 5 Advantages and benefits of the control logic 500 described. For example, Figure 7 700 is a graph of engine speed versus time, where time is plotted on the X-axis and engine speed is plotted on the Y-axis. The engine speed may be between 0% speed and 100% speed. The engine speed may be the rotational speed of the HP shaft 128 and / or the LP shaft 130. Figure 8 is a gap response graph 800 of gap versus time, where time is plotted on the X-axis and gap (e.g., Figure 3 The clearance between the blade tips 192a, 192b and the shrouds 190a, 190b is shown plotted on the Y-axis. The clearance may be between a low clearance value or a small clearance value (e.g., the tip of the rotor blade is close to or in contact with the casing) and a high clearance value or a large clearance value (such that the tip of the rotor blade is away from the casing).

[0078] The graph 700 may include an engine speed line 708 representing the rotational speed of the HP shaft 128 and / or the LP shaft 130 at the particular point in time shown. Conversely, the clearance response graph 800 may include an actual clearance line 808 representing actual clearance (e.g., measured, sensed, and / or calculated clearance between the blade tips 192a, 192b and the shrouds 190a, 190b at a particular point in time) and a target clearance line 810 representing target clearance (e.g., clearance generated by the engine controller 402 for optimal performance of the gas turbine engine 100). Figure 8 As shown, the points of the actual clearance line 808 are drawn with circles, and the points of the target clearance line 810 are drawn with stars.

[0079] Should be observed together Figure 7 and Figure 8 , to understand the clearance response of the gas turbine engine 100 resulting from the adjustment of the acceleration rate. For example, each of the graphs 700, 800 shares a common time X-axis and a common number of data points plotted in 10 second increments. Figure 7 and Figure 8 As shown, lines 708, 808 may include initial points 702, 802 associated with a first flight condition 712 (such as a steady-state condition) of the gas turbine engine 100. For example, the initial point 702 may be the engine speed (e.g., the rotational speed of the HP shaft 128 and / or the LP shaft 130) at the first flight condition 712. Similarly, the initial point 802 may be the actual clearance between the turbine blades and the turbine casing at the first flight condition 712. In addition, as shown in FIG. Figure 7 and Figure 8 As shown, the lines 708, 808 may include final points 706, 806 associated with a final flight condition 714 (such as a step climb condition) that is different from the first flight condition 712. For example, the final point 706 may be the engine speed (e.g., the rotational speed of the HP shaft 128 and / or the LP shaft 130) at the final flight condition 714. Similarly, the final point 806 may be a gap at the final flight condition 714. In addition, when the gas turbine engine 100 transitions between the first flight condition 712 and the final flight condition 714, each of the lines 708, 808 may include a plurality of transient points 704, 804 between the initial point 702, 802 and the final point 706, 806.

[0080] The engine controller 402 may determine a plurality of transient target clearances 820 between the initial clearance 816 (e.g., when the gas turbine is operating at the first flight condition 712) and the final target clearance 818 (e.g., when the gas turbine engine is operating at the final flight condition 714). For example, when the gas turbine engine 100 is operating at the first flight condition 712 (e.g., a steady-state condition) and the engine controller 402 receives a request for a final flight condition 714 that is different from the first flight condition 712 (such as a steady-state condition), the engine controller 402 may generate, determine, or set the final target clearance 818 associated with the final flight condition 714 and the plurality of transient target clearances 820 between the initial clearance 816 and the final target clearance 818. The plurality of transient target clearances 816 may be predetermined or may be determined based on current operating conditions of the gas turbine engine 100. For example, in some embodiments, each of the intermediate target clearances may be determined based on a current acceleration rate of the gas turbine engine 100 and / or based on other parameters associated with the gas turbine engine 100. For example, each of the transient target clearances may be determined based on various speeds, accelerations, pressures, and / or temperatures associated with the gas turbine engine 100 .

[0081] When gas turbine engine 100 transitions between first flight condition 712 and final flight condition 714, engine controller 402 may periodically (e.g., as shown in FIG. 1 ) Figure 7 and Figure 8 808 is compared to the target clearance line 810 every 10 seconds (as shown). If the actual clearance value is outside the predetermined margin of the target clearance value at the specific time of the comparison, the engine controller 402 can adjust the acceleration rate of the gas turbine engine 100. For example, Figure 7 and Figure 8 As shown together, during the transient state between the first flight condition 712 and the final flight condition 714, the engine controller 402 may adjust the acceleration rate (e.g., as determined by Figure 7 708 in FIG. 7 ).

[0082] Fig. 9 and 10 The above reference is graphically described Figure 6 Advantages and benefits of the control logic 600 described. For example, Fig. 9 900 is a graph of engine speed versus time, where time is plotted on the X-axis and engine speed is plotted on the Y-axis. The engine speed may be between 0% speed and 100% speed. The engine speed may be the rotational speed of the HP shaft 128 and / or the LP shaft 130. Fig.101000 is a gap response curve graph of gap versus time, where time is plotted on the X-axis and gap (e.g., gap between blade tip 192a, 192b and shroud 190a, 190b) is plotted on the Y-axis. The gap can be between a low or small gap value (e.g., close to 0% with the tip of the rotor blade close to or in contact with the casing) and a high or large gap value (such that the tip of the rotor blade is away from the casing).

[0083] The graph 900 may include an engine speed line 908 that represents the rotational speed of the HP shaft 128 and / or the LP shaft 130 at the particular point in time shown. Conversely, the clearance response curve 1000 may include an actual clearance line 1008 that represents actual clearance (e.g., measured, sensed, and / or calculated clearance at a particular point in time) and a target clearance line 1010 that represents target clearance (e.g., clearance generated by the engine controller 402 for optimal performance of the gas turbine engine 100). Fig.10 As shown, points of the actual clearance line 1008 are drawn with circles, and points of the target clearance line 1010 are drawn with stars.

[0084] Should be observed together Fig. 9 and Fig.10 , to understand the clearance response of the gas turbine engine 100 resulting from the adjustment of the acceleration rate. For example, each of the graphs 900, 1000 shares a common time X-axis from 0 to 120 seconds. Fig. 9 and Fig.10 As shown, lines 908, 1008 may include initial points 902, 1002 associated with a first flight condition 912 (such as a steady-state condition) of the gas turbine engine 100. For example, the initial point 902 may be the engine speed (e.g., the rotational speed of the HP shaft 128 and / or the LP shaft 130) at the first flight condition 912. Similarly, the initial point 1002 may be the actual clearance between the turbine blades and the shroud at the first flight condition 912. In addition, as shown in FIG. Fig. 9 and Fig.10 As shown, lines 908, 1008 may include final points 906, 1006 associated with final flight condition 914. For example, final point 906 may be the engine speed (e.g., the rotational speed of HP shaft 128 and / or LP shaft 130) at final flight condition 914. Similarly, final point 1006 may be the clearance between turbine blades and turbine casing at final flight condition 914.

[0085] exist Fig. 9 and Fig.10In the illustrated embodiment, the first flight condition 912 may be the same as the final flight condition 914. In addition, as shown, the lines 908, 1008 may include intermediate points 903, 1003 associated with an intermediate flight condition 913, which may be different from the first flight condition 912 and the final flight condition 914. For example, the intermediate point 903 may be the engine speed (e.g., the rotational speed of the HP shaft 128 and / or the LP shaft 130) at the intermediate flight condition 913. Similarly, the intermediate point 1003 may be the actual clearance between the turbine blades and the turbine casing at the intermediate flight condition 913.

[0086] In addition, when the gas turbine engine 100 transitions from the first flight condition 912 to the intermediate flight condition 913, and then transitions from the intermediate flight condition 913 to the final flight condition 914, each of the lines 908, 1008 may include a plurality of transient points 904, 1004 between the initial point 902, 1002 and the intermediate point 903, 1003 and between the intermediate point 903, 1003 and the final point 906, 1006.

[0087] The intermediate flight condition 913 may require the gas turbine engine to decelerate (e.g., reduce the rotational speed of the HP shaft 128 and / or the LP shaft 130). Subsequently, the gas turbine engine 100 may be required to reaccelerate to the final flight condition 914. In doing so, the gas turbine engine 100 may encounter a hot rotor reacceleration condition. For example, during deceleration, the casing cools and thermally contracts, while the rotor (e.g., the shaft and rotor disks) remains hot and thermally expands due to the mass difference between the casing and the rotor. Therefore, when the gas turbine engine reaccelerates (e.g., between the intermediate flight condition 913 and the final flight condition 914), the turbine blades have less room for mechanical expansion because the rotor is still hot and thermally expands, and the casing is cooled and thermally contracts.

[0088] The engine controller 402 can account for any hot rotor re-acceleration conditions by linking the deceleration rate to the transient target clearances. For example, the engine controller 402 can determine a plurality of transient target clearances 1024 between the initial target clearance 1016 and the intermediate target clearance 1022. When the gas turbine engine 100 is operating at the first flight condition 912 and the engine controller 402 receives a demand for the intermediate flight condition 913 (e.g., deceleration), the engine controller 402 can generate, determine, or set the intermediate target clearance 1022 associated with the intermediate flight condition 913 and a plurality of transient target clearances 1024 between the initial target clearance 1016 and the intermediate target clearance 1022. This advantageously allows tighter clearances to be achieved when the gas turbine engine transitions between the first flight condition and the second flight condition.

[0089] Similarly, when the gas turbine engine 100 is operating at the intermediate flight condition 913, and the engine controller 402 receives a demand for the final flight condition 914 (e.g., requiring hot rotor reacceleration of the gas turbine engine 100), the engine controller 402 may generate, determine, or set the final target clearance 1026 associated with the final flight condition 914 without any transient target clearances between the intermediate target clearance 1022 and the final target clearance 1026. That is, the engine controller 402 may determine that hot rotor reacceleration is required, and may set the final target clearance 1026 without any transient target clearances, which advantageously allows mechanical expansion of the turbine blades with sufficient clearance or radial space without causing a pinch or rubbing event with the casing.

[0090] When gas turbine engine 100 transitions between first flight condition 912 and intermediate flight condition 913, and then transitions between intermediate flight condition 913 and final flight condition 914, engine controller 402 may periodically (e.g., as shown in FIG. 1 ) perform a Fig. 9 and Fig.10 10 seconds) the actual clearance line 1008 is compared to the target clearance line 1010. If the actual clearance value is outside the predetermined margin of the target clearance value at the specific time point of the comparison, the engine controller 402 can adjust the deceleration rate or acceleration rate of the gas turbine engine 100. For example, Fig. 9 and Fig.10 As shown together, during the transition state between the first flight condition 912 and the intermediate flight condition 914, the engine controller 402 may adjust the acceleration rate (as determined by the Fig. 9 , as shown by the changing slope of the engine speed line 908 in FIG.

[0091] Additionally or alternatively, the acceleration / deceleration rate may be adjusted based on other parameters, such as exhaust temperature, accelerometer data, strain measurements (e.g., via one or more strain gauges). For example, the acceleration may be adjusted based on the exhaust temperature (which may be measured with one or more sensors, or which may be calculated). Specifically, when the exhaust temperature exceeds a first predetermined value, or drops below a second predetermined value, the acceleration / deceleration may be adjusted. Similarly, if the gas turbine is experiencing excessive strain, the acceleration / deceleration may be modified (e.g., the acceleration may be reduced).

[0092] Fig.11 A flow chart of a method 1100 of operating a gas turbine engine, such as the gas turbine engine 100 discussed above, is provided. For example, some or all of the method 1100 may be performed by the engine controller 402 ( Figure 4). The gas turbine engine may include a first component (such as casing 198) and a second component (such as rotor blade 166) rotatable relative to the first component. A gap is defined between the first component and the second component (e.g., directly between the first component and the second component).

[0093] In many embodiments, the method may include operating the gas turbine engine at a first flight condition at (1102). In an exemplary embodiment, the first flight condition may be a steady-state cruise condition in which the gas turbine engine is neither accelerating nor decelerating. For example, in a steady-state cruise condition, the gas turbine engine (e.g., the HP shaft and / or the LP shaft) may rotate at a substantially constant speed. In other embodiments, the first flight condition may be an acceleration condition in which the gas turbine engine is increasing speed. Alternatively, the first flight condition may be a deceleration condition in which the gas turbine engine is decreasing speed.

[0094] In various embodiments, method 1100 may include receiving a request for a second flight condition that is different from the first flight condition at (1104). In an exemplary embodiment, the second flight condition may be a step climb condition. A step climb condition occurs when a pilot of an aircraft chooses to increase the altitude at which the aircraft is traveling. According to current FAA regulations, altitude steps typically occur in 2,000-foot increments. This means, for example, that a pilot of an aircraft flying at 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, the pilot modifies the control of the autopilot / autothrottle system of the flight control system 420 to request that the aircraft ascend to a desired cruising altitude. The flight control system then uses a predetermined algorithm to increase engine power in order to cause the aircraft to climb. The engine controller may receive this request by monitoring the position of the throttle and observing a specific change in the angle of the throttle, which indicates a request for a change in engine RPM or a step climb condition. The request may also be received by the autopilot / autothrottle control system of the flight control system, and the pilot may request an increase from one specific altitude to a second specific altitude. In response to the request, the flight control system sends a signal to the engine controller requesting an increase in engine power, which causes the gas turbine engine to accelerate.

[0095] In other embodiments, the second flight condition may be an acceleration condition in which the gas turbine engine is increasing speed. Alternatively, the second flight condition may be a deceleration condition in which the gas turbine engine is decreasing speed.

[0096] according to Fig.11In an exemplary embodiment of the present invention, engine control system 402 may receive a request to increase engine power and determine whether the aircraft is currently operating in a steady-state cruise condition. For example, to determine whether the aircraft is in a cruise condition, engine control system 402 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. In response to determining that a cruise condition exists, engine control system 402 then interprets the request to increase engine power as a request for a step climb condition.

[0097] In an exemplary embodiment, method 1100 may include determining a final target clearance and a transient target clearance between the first component and the second component at (1106). The final target clearance may be associated with the second flight condition. That is, the final target clearance may be an optimized clearance when the gas turbine engine reaches the second flight condition. The transient target clearance may be generated by an engine controller, and the transient target clearance may be an optimized clearance for the gas turbine engine when the gas turbine engine transitions between the first flight condition and the second flight condition.

[0098] like Fig.11 As shown, in many embodiments, method 1100 may include adjusting the engine acceleration rate to a nominal acceleration rate at (1108). The nominal acceleration rate may be a predetermined acceleration rate initiated by default by the engine controller when the gas turbine engine is required to accelerate / decelerate. In embodiments where the gas turbine engine is decelerating between a first flight condition and a second flight condition, the nominal acceleration rate may be negative. The nominal acceleration rate may be an angular acceleration value (i.e., a change in angular velocity RPM per unit time), such as revolutions per minute / second (e.g., RPM / s).

[0099] In an exemplary embodiment, the method 1100 may include comparing the actual clearance to the transient target clearance after the time increment at (1110) (e.g., using an engine controller). The actual clearance may be received and / or calculated by the engine controller. For example, in some embodiments, the method may include receiving data indicating an actual clearance between the first component and the second component. The actual clearance is at least one of a measured clearance captured by a sensor and a calculated clearance specific to the gas turbine engine at that point in time. For example, the actual clearance may be indicated by the sensor and expressed in units of distance (such as millimeters).

[0100] In some embodiments, the comparison at (1110) may further include determining at (1112) whether the actual gap is within a predetermined margin of the transient target gap. The predetermined margin may be a ±15% margin of the transient target gap, or such as a ±10% margin of the transient target gap, or such as a ±5% margin of the transient target gap, or such as a ±1% margin of the transient target gap. When the actual gap is neither greater than a maximum margin threshold nor less than a minimum margin threshold, the actual gap is within the predetermined margin. Similarly, when the actual gap exceeds the maximum margin threshold or is less than the minimum margin threshold, the actual gap is outside the predetermined margin. For example, if the actual gap is 1 mm and the predetermined margin of the transient target gap is from 0.5 mm (minimum margin threshold) to 1.5 mm (maximum margin threshold), the actual gap is within the predetermined margin of the transient target gap. Conversely, for example, if the actual clearance is 2 mm, and the predetermined margin of the transient target clearance is from 0.5 mm (minimum margin threshold) to 1.5 mm (maximum margin threshold), the actual clearance is outside the predetermined margin of the transient target clearance. In this case, the actual clearance is too large, and the engine acceleration may increase to cause more mechanical / thermal growth of the blades, thereby reducing the actual clearance to within the predetermined margin of the transient target clearance.

[0101] In certain embodiments, the method 1100 may include adjusting the engine acceleration rate based on a comparison between the actual clearance and the transient target clearance at (1114). As a result, the transient target clearance may be adjusted based at least in part on the adjustment to the engine acceleration rate. For example, when the actual clearance is outside a predetermined margin of the transient target clearance (e.g., exceeds a maximum margin threshold or is below a minimum margin threshold), the engine acceleration rate may be adjusted (e.g., increased or decreased). For example, the engine acceleration rate may be increased / decreased by about 1% to about 90%, or such as about 1% to about 60%, or such as about 1% to about 30%, or such as about 1% to about 15%. In particular, if the actual clearance exceeds a maximum margin threshold of a predetermined margin (i.e., the clearance is too large, which may result in inefficiency), the engine acceleration rate may be increased. Conversely, if the actual clearance is below a minimum margin threshold of a predetermined margin (e.g., the clearance is too small, which may result in a squeeze / friction event), the engine acceleration rate may be decreased.

[0102] When the gas turbine engine transitions between the first flight condition and the second flight condition, the transient target clearance can be adjusted in real time based on adjustments to other parameters of the gas turbine engine. Specifically, the transient target clearance can be adjusted each time the engine acceleration rate is adjusted. For example, if the acceleration rate of the gas turbine engine is adjusted based on a comparison between the actual clearance and the transient target clearance, the transient target clearance can also be adjusted. In other words, the transient target clearance can be a function of the engine speed (e.g., the core speed or the rotational speed of the HP shaft and / or the LP shaft), so that when the gas turbine engine accelerates (or decelerates), the transient clearance target will be adjusted based on the engine speed.

[0103] In some embodiments, the gas turbine engine includes a clearance adjustment system configured to adjust a clearance between a first component and a second component. In such embodiments, the method 1100 may include, at (1116) (which is an optional step as shown in a dashed box), causing the clearance adjustment system to adjust the clearance based on a comparison between the actual clearance and the transient target clearance. For example, in addition to adjusting the acceleration of the gas turbine engine, or as an alternative to adjusting the acceleration of the gas turbine engine, when the actual clearance is outside a predetermined margin of the transient target clearance (e.g., exceeding a maximum margin threshold or being below a minimum margin threshold), the clearance adjustment system may make an adjustment (e.g., increase or decrease).

[0104] In many embodiments, steps 1110 to 1114 may be repeated until the actual clearance is within a predetermined acceptable range of the final target clearance and / or when the actual clearance is equal to the final target clearance. Alternatively or additionally, steps 1110 to 1114 may be repeated until the gas turbine engine has achieved a second flight condition (or is operating under a second flight condition). For example, method 1100 may include repeating the steps of: comparing between the actual clearance and the transient target clearance, and adjusting the engine acceleration rate based on the comparison after the time increment until the actual clearance is within a predetermined acceptable range of the final target clearance. The predetermined acceptable range may be between approximately ±15% of the final target clearance, or such as between approximately ±10% of the final target clearance, or such as between approximately ±5% of the final target clearance, or such as between approximately ±1% of the final target clearance.

[0105] In various embodiments, method 1100 may include determining whether engine operating conditions allow modification of the nominal acceleration rate. For example, method 1100 may include determining whether a gas turbine engine operating condition for which an optimized acceleration rate (e.g., a deviation or modification from the nominal acceleration rate) is marked. Many operating conditions of a gas turbine engine for which an optimized acceleration rate may be marked, such as a maximum acceleration condition (when the gas turbine engine accelerates from a first speed to a second speed), a step climb condition, or others. However, under some operating conditions, it may be optimal to proceed at a nominal acceleration rate. For example, if the aircraft is very close to a runway (such as a takeoff or landing condition). The engine controller may identify when the operating conditions of the gas turbine engine for applying the optimized acceleration rate are optimal to ensure that the optimized acceleration rate is not applied at an inappropriate time. The optimized acceleration rate may refer to a modification of the nominal acceleration rate based on a comparison between a transient target clearance and an actual clearance.

[0106] Fig.12 A block diagram of an engine controller 402 according to an example embodiment of the present disclosure is provided. As shown, the engine controller 402 may include one or more processors 211 and one or more memory devices 212. The one or more processors 211 may include any suitable processing device, such as a microprocessor, a microcontroller, an integrated circuit, a logic device, and / or other suitable processing device. The one or more memory devices 212 may include one or more computer executable or computer readable media, including but not limited to non-transitory computer readable media, RAM, ROM, hard disk drives, flash drives, and / or other memory devices.

[0107] The one or more memory devices 212 may store information accessible by the one or more processors 211, including computer-readable or computer-executable instructions 213 that may be executed by the one or more processors 211. The instructions 213 may include any set of instructions that, when executed by the one or more processors 211, cause the one or more processors 211 to perform an operation. The instructions 213 may be software written in any suitable programming language or may be implemented in hardware. Additionally and / or alternatively, the instructions 213 may be executed in logically and / or virtually separate threads on the processor 211. The memory device 212 may further store data 214 that may be accessed by the processor 211. For example, the data 214 may include models, lookup tables, databases, etc. The data 214 may include Figure 4 Sensor data 415.

[0108] The engine controller 402 may also include a network interface 215 for communicating, for example, with other devices communicatively coupled thereto (e.g., via a communication network). The network interface 215 may include any suitable components for interfacing with one or more networks, including, for example, transmitters, receivers, ports, controllers, antennas, and / or other suitable components. One or more devices may be configured to receive one or more commands, control signals, and / or data from the engine controller 402, or to provide one or more commands, control signals, and / or data to the engine controller 402.

[0109] The technology discussed herein refers to computer-based systems, actions taken by computer-based systems, information sent to computer-based systems, and information from computer-based systems. It should be understood that the inherent flexibility of computer-based systems allows for multiple possible configurations, combinations, and divisions of tasks and functions between and among components. For example, the processing discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and applications can be implemented on a single system or distributed on multiple systems. Distributed components can operate sequentially or in parallel.

[0110] The systems and methods disclosed herein facilitate continuous communication between the acceleration rate and target clearance of a gas turbine engine, which advantageously prevents the rotor blades from pinching against one or more stationary components (such as a casing). At one or more points during acceleration or deceleration, the system can determine whether the gas turbine engine is on pace to achieve the final target clearance, and the system can adjust the acceleration / deceleration as needed to achieve the final clearance without causing a pinch event. This advantageously extends the life of the components of the gas turbine engine by preventing damage and allows tighter clearances, thereby improving efficiency.

[0111] Further aspects are provided by the subject matter of the following clauses:

[0112] A gas turbine engine comprising: a first component; a second component, the second component being rotatable relative to the first component, wherein a clearance is defined between the first component and the second component; and an engine controller having one or more processors, the one or more processors being configured to: operate the gas turbine engine at a first flight condition; receive a request for a second flight condition different from the first flight condition; determine a final target clearance and a transient target clearance between the first component and the second component, the final target clearance being associated with the second flight condition; adjust an engine acceleration rate to a nominal acceleration rate; compare an actual clearance to the transient target clearance after a time increment; and adjust the engine acceleration rate based on the comparison between the actual clearance and the transient target clearance, wherein the transient target clearance is adjusted based at least in part on the adjustment to the engine acceleration rate.

[0113] A gas turbine engine according to any preceding clause, wherein, when comparing the actual clearance with the transient target clearance, one or more processors are further configured to: determine whether the actual clearance is within a predetermined margin of the transient target clearance; and when the actual clearance is not within the predetermined margin of the transient target clearance, adjust the engine acceleration rate.

[0114] A gas turbine engine according to any preceding clause, wherein, when comparing the actual clearance with the transient target clearance, one or more processors are further configured to: determine that the actual clearance exceeds the transient target clearance by a predetermined margin; and increase engine acceleration until the actual clearance is within the predetermined margin of the transient target clearance.

[0115] A gas turbine engine according to any preceding clause, wherein, when comparing the actual clearance with the transient target clearance, one or more processors are further configured to: determine that the actual clearance is less than a predetermined margin of the transient target clearance; and reduce engine acceleration until the actual clearance is within the predetermined margin of the transient target clearance.

[0116] A gas turbine engine according to any preceding clause, wherein said one or more processors are further configured to repeat said comparison between said actual clearance and said transient target clearance and said adjustment of said engine acceleration rate based on said comparison after said time increment until said actual clearance is within a predetermined acceptable range of said final target clearance.

[0117] A gas turbine engine according to any preceding clause, wherein the one or more processors are further configured to: repeat the comparison between the actual clearance and the transient target clearance and the adjustment of the engine acceleration rate based on the comparison after the time increment until the gas turbine engine is operating under the second flight condition.

[0118] A gas turbine engine according to any preceding clause, wherein the one or more processors are further configured to receive data indicating the actual clearance between the first component and the second component, the actual clearance being at least one of a measured clearance captured by a sensor and a calculated clearance specific to the gas turbine engine at that point in time.

[0119] The gas turbine engine according to any preceding clause, further comprising a clearance adjustment system, the clearance adjustment system being configured to adjust the clearance between the first component and the second component, and wherein the one or more processors are further configured to cause the clearance adjustment system to adjust the clearance based on the comparison between the actual clearance and the transient target clearance.

[0120] A gas turbine engine as claimed in any preceding clause, wherein the transient target clearance is adjusted each time the engine acceleration rate is adjusted.

[0121] A gas turbine engine as claimed in any preceding clause, wherein the one or more processors are further configured to determine whether engine operating conditions warrant modification of the nominal acceleration rate.

[0122] A gas turbine engine as described in any preceding clause, wherein the first flight condition comprises a steady-state cruise condition, and wherein the second flight condition comprises a step-climb condition.

[0123] A non-transitory computer-readable medium including computer-executable instructions that, when executed by one or more processors of a computing system associated with a gas turbine engine, cause the one or more processors to: operate the gas turbine engine at a first flight condition; receive data indicating an actual clearance between a first component of the gas turbine engine and a second component of the gas turbine engine, the second component being rotatable relative to the first component, the actual clearance being at least one of a measured clearance captured by a sensor and a calculated clearance specific to the gas turbine engine at that point in time; receive a demand for a second flight condition different from the first flight condition; determine a final target clearance and a transient target clearance between the first component and the second component, the final target clearance being associated with the second flight condition; adjust an engine acceleration rate to a nominal acceleration rate; compare the actual clearance to the transient target clearance after a time increment; and adjust the engine acceleration rate based on the comparison between the actual clearance and the transient target clearance, wherein the transient target clearance is adjusted based at least in part on the adjustment to the engine acceleration rate.

[0124] A non-transitory computer-readable medium as described in any preceding clause, wherein, when comparing the actual gap with the transient target gap, the one or more processors are configured to: determine whether the actual gap is within a predetermined margin of the transient target gap; and when the actual gap is not within the predetermined margin of the transient target gap, adjust the engine acceleration rate.

[0125] A non-transitory computer-readable medium as described in any preceding clause, wherein the one or more processors are further configured to repeat the steps of: comparing between the actual clearance and the transient target clearance; and adjusting the engine acceleration rate based on the comparison after the time increment until the actual clearance is within a predetermined acceptable range of the final target clearance.

[0126] A non-transitory computer-readable medium as described in any preceding clause, wherein the one or more processors are further configured to repeat the steps of: comparing between the actual clearance and the transient target clearance; and adjusting the engine acceleration rate based on the comparison after the time increment until the gas turbine engine operates at the second flight condition.

[0127] A non-transitory computer-readable medium according to any preceding clause, wherein the gas turbine engine includes a clearance adjustment system, the clearance adjustment system being configured to adjust the clearance between the first component and the second component, and one or more processors being configured to cause the clearance adjustment system to adjust the clearance based on a comparison between the actual clearance and the transient target clearance.

[0128] The non-transitory computer readable medium of any preceding clause, wherein the transient target clearance is adjusted each time the engine acceleration rate is adjusted.

[0129] The non-transitory computer readable medium of any preceding clause, wherein the one or more processors are further configured to determine whether engine operating conditions warrant modification of the nominal acceleration rate.

[0130] The non-transitory computer-readable medium of any preceding clause, wherein operating the gas turbine engine at a first flight condition comprises operating the gas turbine engine at a steady-state cruise condition, and wherein the second flight condition is a step-climb condition.

[0131] A method of operating a gas turbine engine, the gas turbine engine comprising a first component and a second component rotatable relative to the first component, a clearance defined between the first component and the second component, the method comprising: operating the gas turbine engine at a first flight condition; receiving a request for a second flight condition different from the first flight condition; determining a final target clearance and a transient target clearance between the first component and the second component, the final target clearance being associated with the second flight condition; adjusting an engine acceleration rate to a nominal acceleration rate; comparing an actual clearance to the transient target clearance after a time increment; and adjusting the engine acceleration rate based on the comparison between the actual clearance and the transient target clearance, wherein the transient target clearance is adjusted based at least in part on the adjustment to the engine acceleration rate.

[0132] A method according to any preceding clause, wherein comparing the actual clearance with the transient target clearance further comprises: determining whether the actual clearance is within a predetermined margin of the transient target clearance; and adjusting the engine acceleration rate when the actual clearance is not within the predetermined margin of the transient target clearance.

[0133] A method according to any preceding clause, further comprising repeating the steps of: comparing between said actual clearance and said transient target clearance; and adjusting said engine acceleration rate based on the comparison after said time increment until said actual clearance is within a predetermined acceptable range of said final target clearance.

[0134] The method according to any preceding clause, further comprising repeating the steps of: comparing between the actual clearance and the transient target clearance; and adjusting the engine acceleration rate based on the comparison after the time increment until the gas turbine engine is operating at the second flight condition.

[0135] The method of any preceding clause, further comprising receiving data indicative of the actual clearance between the first component and the second component, the actual clearance being at least one of a measured clearance captured by a sensor and a calculated clearance specific to the gas turbine engine at that point in time.

[0136] A method according to any preceding clause, wherein the gas turbine engine includes a clearance adjustment system configured to adjust the clearance between the first component and the second component, and wherein the method further comprises: causing the clearance adjustment system to adjust the clearance based on the comparison between the actual clearance and the transient target clearance.

[0137] A method as described in any preceding clause, wherein the transient target clearance is adjusted each time the engine acceleration rate is adjusted.

[0138] A method as in any preceding clause, further comprising determining whether engine operating conditions permit modification of said nominal acceleration rate.

[0139] A method as in any preceding clause, wherein operating the gas turbine engine in a first flight condition comprises operating the gas turbine engine in a steady-state cruise condition, and wherein the second flight condition is a step-climb condition.

[0140] This written description uses examples to disclose the present disclosure, including the best mode, and also enables any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined methods. The patent scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, these other examples are intended to fall within the scope of the claims.

Claims

1. A gas turbine engine, characterized in that: include: first component; a second component rotatable relative to the first component, wherein a gap is defined between the first component and the second component; as well as An engine controller having one or more processors configured to: operating the gas turbine engine under a first flight condition; receiving a request for a second flight condition different from the first flight condition; determining a final target clearance and a transient target clearance between the first component and the second component, the final target clearance being associated with the second flight condition; Adjust the engine acceleration rate to the nominal acceleration rate; comparing the actual gap to the instantaneous target gap after the time increment; as well as The engine acceleration rate is adjusted based on a comparison between the actual lash and the transient target lash, wherein the transient target lash is adjusted based at least in part on the adjustment to the engine acceleration rate.

2. The gas turbine engine according to claim 1, characterized in that in, When comparing the actual gap to the transient target gap, the one or more processors are further configured to: determining whether the actual gap is within a predetermined margin of the transient target gap; and When the actual lash is not within the predetermined margin of the transient target lash, the engine acceleration rate is adjusted.

3. The gas turbine engine according to claim 1, characterized in that in, When comparing the actual gap to the transient target gap, the one or more processors are further configured to: determining that the actual gap exceeds the transient target gap by a predetermined margin; and Engine acceleration is increased until the actual lash is within the predetermined margin of the transient target lash.

4. The gas turbine engine according to claim 1, characterized in that in, When comparing the actual gap to the transient target gap, the one or more processors are further configured to: determining that the actual gap is less than the transient target gap by a predetermined margin; and Engine acceleration is reduced until the actual lash is within the predetermined margin of the transient target lash.

5. The gas turbine engine according to claim 1, characterized in that in, The one or more processors are further configured to: The comparing of the actual clearance with the transient target clearance and the adjusting of the engine acceleration rate based on the comparing after the time increment are repeated until the actual clearance is within a predetermined acceptable range of the final target clearance.

6. The gas turbine engine according to claim 1, characterized in that in, The one or more processors are further configured to: The comparing of the actual clearance with the transient target clearance and the adjusting of the engine acceleration rate based on the comparing after the time increment are repeated until the gas turbine engine is operating at the second flight condition.

7. The gas turbine engine according to claim 1, characterized in that in, The one or more processors are further configured to: Data indicative of the actual clearance between the first component and the second component is received, the actual clearance being at least one of a measured clearance captured by a sensor and a calculated clearance specific to the gas turbine engine at that point in time.

8. The gas turbine engine according to claim 1, characterized in that further comprising a gap adjustment system configured to adjust the gap between the first component and the second component, and wherein the one or more processors are further configured to: The gap adjustment system is caused to adjust the gap based on the comparison between the actual gap and the transient target gap.

9. The gas turbine engine according to claim 1, characterized in that in, The transient target clearance is adjusted each time the engine acceleration rate is adjusted.

10. The gas turbine engine according to claim 1, characterized in that in, The one or more processors are further configured to: A determination is made as to whether engine operating conditions warrant modification of the nominal acceleration rate.

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