Systems and methods for detecting and responding to icing conditions in a gas turbine engine

Abstract

An engine controller may include a processor and a feedback interface electrically coupled to the processor and one or more electric motors and generators deployed in an aircraft engine. A memory electrically coupled to the processor is configured to store reference data indicating the presence of an icing condition. The processor is configured to monitor feedback signals received from the one or more electric motors and generators at the feedback interface and, based on a correspondence between the feedback signals and the reference data, identify whether the feedback signals indicate the presence of an icing condition in the aircraft engine. The processor is configured to initiate an icing condition mitigation action in response to identifying the presence of an icing condition.

Description

Technical Field

[0001] These teachings relate generally to engines, and more particularly to systems and methods for detecting and responding to icing conditions in gas turbine jet engines. Background Art

[0002] Gas turbine and similar jet engines may encounter icing conditions during operation. These icing conditions may include ice accumulation on the fan or other turbine engine components, as well as the shedding of previously accumulated ice. Such icing and shedding events can affect turbine engine operation by increasing engine loads, creating unstable load imbalances, producing sudden and unexpected changes in engine loads, and / or increasing the risk of stalling the aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] Various needs are met, at least in part, by providing systems and methods for detecting and responding to icing conditions in a gas turbine engine as described in the following detailed description, particularly when studied in conjunction with the accompanying drawings. A full and enabling disclosure of various aspects of the present description, 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, wherein:

[0004] Figure 1 is a cross-sectional view of a gas turbine engine for an aircraft;

[0005] Figure 2 is a schematic diagram of a gas turbine engine control system according to some embodiments;

[0006] Figure 3 is a graph of the turbine engine operating line under various conditions described herein;

[0007] Figure 4 is a graph of a gas turbine engine parameter versus time in response to the presence of icing conditions; and

[0008] Figure 5 is a flow chart of a method for detecting and responding to icing conditions on an aircraft engine, according to some embodiments.

[0009] The elements shown in the figures are shown for simplicity and clarity and are not necessarily drawn to scale. For example, the size and / or relative positioning of certain elements in the figures may be exaggerated relative to other elements to help improve understanding of the various embodiments of the present teachings. In addition, common but easy-to-understand elements that are useful or necessary in commercially feasible embodiments are generally not depicted to facilitate a clearer view of these various embodiments of the present teachings. Certain actions and / or steps may be described or depicted in a specific order of occurrence, and those skilled in the art will understand that such specificity about the order is not actually required. DETAILED DESCRIPTION

[0010] The terms and expressions used herein have the ordinary technical meanings assigned to them by those skilled in the art as described above, unless otherwise specifically defined herein. Unless expressly stated otherwise, the word "or" as used herein should be interpreted as having a disjunctive rather than a conjunctive structure. Unless otherwise specified herein, the terms "coupled," "fixed," "attached to," and the like refer to directly coupled, fixed, or attached, as well as indirectly coupled, fixed, or attached through one or more intermediate components or features.

[0011] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0012] As used herein throughout the specification and claims, approximating language is used to modify any quantitative representation that can be permissibly varied without resulting in a change in the basic function to which it is related. Thus, a value modified by one or more terms (e.g., "about," "approximately," and "substantially") is not limited to the precise value specified. In at least some cases, approximating language may correspond to the precision of an instrument used to measure the value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximating language may mean within a 10% margin.

[0013] Current engine control systems utilize specialized sensors to identify the presence of icing conditions and typically employ electric or pneumatic anti-ice systems to melt the accumulated ice to restore engine performance and prevent uncontrolled shedding events. However, the additional sensors and / or de-icing systems employed by these systems can be expensive, degrade in performance over time, add unnecessary complexity to the engine, and employ complex control procedures to protect the hardware from stalling in the event of ice shedding and ice ingestion into the core. Furthermore, aircraft utilizing current engine de-icing systems can generate higher rpm (e.g., higher N1 values) on the low-pressure spool relative to the maximum normal operating rpm until the de-icing system is activated. These are significant challenges in aviation applications.

[0014] In general, various aspects of the present disclosure relate to an engine controller having a feedback interface electrically coupled to one or more motor-generators deployed in an aircraft engine for providing icing protection for the fan, mid-fan, and supercharger sections of the aircraft engine. Generally, the engine controller described herein is configured to detect or identify the presence of an icing condition (e.g., an accretion event or a de-icing event) on the aircraft engine and employ various regulatory or mitigation procedures in response to the detected icing condition. Signals received by the engine controller from the motor-generators and / or other torque sensors deployed on the aircraft engine may be used to identify the presence of an icing condition. In some embodiments, a high-bandwidth portion of a feedback signal from the motor-generators or other aspects of the hybrid electrical architecture of the aircraft engine is used to identify ice accumulation on surfaces that affect the torsional characteristics of the engine shaft.

[0015] In some embodiments, the engine controller utilizes an electric motor-generator to respond to the detection of an icing condition. Specifically, an electric motor-generator coupled to a low-speed or high-speed shaft of the engine can transfer power from a low-pressure shaft or fan shaft to a high-pressure shaft or core shaft to provide increased operability margin (e.g., increase N2 operability margin) before a future de-icing event occurs. The low-speed shaft can include a low-pressure shaft, and the high-speed shaft can include a high-pressure shaft. Additionally or alternatively, the engine controller can employ other methods to increase operability margin, such as changing the variable geometry of one or more engine components, triggering bleed air from the engine compressor, or other similar methods known in the art.

[0016] Additionally, the engine controller can utilize the motor-generator to initiate a de-icing sequence torque resonance mode. The torque resonance mode induces a disturbance in the engine's fan shaft to excite specific shaft modes that can shed ice while simultaneously protecting core operability by injecting power into the core shaft. In some embodiments, the disturbance is induced by periodically varying power extraction from the fan shaft while using steady-state (e.g., non-periodic) N2 power injection to protect core operability.

[0017] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a cross-sectional view of a gas turbine engine. The gas turbine engine is a high-bypass turbofan jet engine, referred to herein as a "turbofan engine 10." The turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 for reference) and a radial direction R. Generally, the turbofan engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14.

[0018] Figure 1The core turbine engine 16 shown includes a generally tubular outer casing 18 defining an annular inlet 20. The tubular outer casing 18 encloses, in serial flow relationship: a compressor section including a low-pressure (LP) compressor 22 and a high-pressure (HP) compressor 24; a combustion section 26; a turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) spool 36 drivingly connects the LP turbine 30 to the LP compressor 22.

[0019] The fan section 14 may include a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 generally in a radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operably coupled to a suitable actuating member 44, which is configured to collectively and uniformly change the pitch of the fan blades 40. The fan blades 40, disk 42, and actuating member 44 are rotatable together about the longitudinal centerline 12 via the LP spool 36 across a power gearbox 46. The power gearbox 46 includes a plurality of gears for reducing the rotational speed of the LP spool 36 to a more efficient fan speed. It should be understood that additional configurations of the fan section 14 are also contemplated, such as those employing non-ducted or non-variable pitch designs.

[0020] Still refer to Figure 1 In the embodiment of the present invention, the disk 42 is covered by a rotatable forward hub 48 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 40. In addition, the fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the variable pitch fan 38 and / or at least a portion of the core turbine engine 16. It should be understood that the outer nacelle 50 is configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the outer nacelle 50 extends above the outer portion of the core turbine engine 16, thereby defining a bypass airflow passage 56 therebetween.

[0021] During operation of turbofan engine 10, a volume of air 58 enters turbofan engine 10 through outer nacelle 50 and associated inlet 60 of fan section 14. As this volume of air 58 passes through fan blades 40, a first portion 62 of air 58 is directed or channeled into bypass airflow passage 56, as indicated by arrows, while a second portion 64 of air 58 is directed or channeled into LP compressor 22, as indicated by arrows. The ratio between first portion 62 of air 58 and second portion 64 of air 58 is generally referred to as the bypass ratio. The pressure of second portion 64 of air 58 is then increased as it is directed through HP compressor 24 and into combustion section 26, where it mixes with fuel and combusts to provide combustion gases 66. Combustion gases 66 are then directed through the hot flow paths, or hot section flow paths, of HP turbine 28 and LP turbine 30, where a portion of the thermal and / or kinetic energy is extracted from combustion gases 66.

[0022] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion 62 of the air 58 is significantly increased as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbofan engine 10, also providing propulsive thrust.

[0023] However, it should be understood that Figure 1 The illustrated turbofan engine 10 is merely an example, and in other embodiments, aspects of the present disclosure may additionally or alternatively be applied to any other suitable gas turbine engine. For example, in other embodiments, the turbofan engine 10 may be any other suitable aviation gas turbine engine, such as a turbojet engine, a turboshaft engine, a turboprop engine, etc. Furthermore, in other embodiments, the turbofan engine 10 may include or be operably connected to any other suitable accessory system. Additionally or alternatively, the turbofan engine 10 may not include or be operably connected to one or more of the accessory systems discussed above.

[0024] Now refer to Figure 2 , showing a method for a gas turbine engine (e.g. Figure 1FIG1 is a schematic diagram of a control system 100 for a turbofan engine 10 of FIG1. The control system 100 includes an engine controller 102, such as a full authority digital engine controller (FADEC), configured to control one or more electric motor generators (e.g., power converters) that are physically geared to portions of the turbofan engine 10. Specifically, the one or more electric motor generators may include a low-pressure electric motor generator (LP-EM) 104 and a high-pressure electric motor generator (HP-EM) 106. The LP-EM 104 is geared to low-pressure components 105 of the turbofan engine 10, including the LP compressor 22, the fan section 14, the LP spool 36, and the LP turbine 30. The HP-EM 106 is geared to high-pressure components 107 of the turbofan engine 10, including the HP turbine 28, the HP compressor 24, and the HP spool 34. Both the LP-EM 104 and the HP-EM 106 are electrically coupled to the engine controller 102 via a data bus 108.

[0025] The engine controller 102 includes a processor 110, a memory 112, a feedback interface 114, a control interface 116, and a sensor interface 118. The processor 110 may include, for example, a microprocessor, a system on a chip, an application-specific integrated circuit (ASIC), a field-programmable gate array (FPGA), etc. The memory 112 may include, for example, a charge-based storage medium (e.g., an electrically erasable programmable read-only memory (EEPROM) or a random access memory (RAM)) or other non-transitory computer-readable medium (e.g., an optical or magnetic storage device). The memory 112 stores machine-readable instructions 120 that are executable by the processor 110 to perform the various methods described herein, and reference data 124 that indicates the presence of an icing condition (e.g., an icing event or a deicing event) on the turbofan engine 10.

[0026] Feedback interface 114 is electrically coupled to data bus 108 and is configured to receive feedback signals 208 from LP-EM 104 and / or HP-EM 106 (see Figure 4 ). Processor 110 monitors feedback signal 208 received at feedback interface 114 and identifies, based on a correspondence between the feedback signal and reference data 124, whether the feedback signal indicates that an icing condition currently exists in turbofan engine 10. Feedback signal 208 serves as an indicator for monitoring high-frequency speed and / or torque oscillations in low-pressure component 105 and / or high-pressure component 107. In particular, processor 110 interprets the high-frequency portion of feedback signal 208 as meaningful information regarding the presence of an icing condition (e.g., ice accretion and / or de-icing event on or from a particular surface of turbofan engine 10).

[0027] In response to identifying the presence of an icing condition, processor 110 is configured to initiate icing condition mitigation actions. Icing condition mitigation actions may help terminate the icing condition and / or reduce the stall probability of turbofan engine 10, which is typically increased by the presence of an icing condition. As described herein, if a de-icing event is detected or anticipated, or if operating in a de-icing sequence torque resonance mode as described herein, icing condition mitigation actions may include timed or coordinated power transfers to high voltage components 107 to improve core operability.

[0028] Control interface 116 is coupled to component 126 of turbofan engine 10, and processor 110 can instruct component 126 to initiate icing condition mitigation actions. In particular, processor 110 can control or modify operating conditions of turbofan engine 10 to initiate icing condition mitigation actions. Sensor interface 118 can electrically connect engine controller 102 to torque sensors 130 and 132, which are respectively gear-coupled to the LP system and HP system of turbofan engine 10. In some embodiments, processor 110 can utilize status signals received from torque sensors 130 and 132 at sensor interface 118 in conjunction with or in lieu of monitoring feedback signals on feedback interface 114 to identify the presence of an icing condition. Torque sensors 130 and 132 can include surface acoustic wave (SAW) sensors or similar sensors. In some embodiments, feedback interface 114, control interface 116, and sensor interface 118 can be part of a unified input and output or communication interface of engine controller 102.

[0029] In some embodiments, components 126 may include LP-EM 104, HP-EM 106, combustion section 26, fuel controller 128, and other components that affect the current operation of turbofan engine 10. For example, current operating conditions may include the pitch of one or more fans (e.g., fan section 14), the speed of one or more cores or shafts (e.g., HP spool 34 and LP spool 36), and / or the overall operating line of turbofan engine 10. As understood by one of ordinary skill in the art, the operating line of turbofan engine 10 defines a desired pressure ratio for turbofan engine 10 for a given steady-state mass flow from turbofan engine 10 (see, e.g., FIG. 1 ). Figure 3 ).

[0030] The presence of an icing condition during operation of turbofan engine 10 may cause changes in the overall performance of turbofan engine 10. To account for these changes, processor 110 is configured to initiate icing condition mitigation actions by instructing component 126 via control interface 116 to change the current operating conditions of turbofan engine 10 in response to processor 110 determining that an icing condition exists. For example, where the operating conditions include the pitch of one or more fans of fan section 14 and the speeds of HP spool 34 and LP spool 36, processor 110 is configured to change the pitch and speeds to maintain the current desired thrust output of the aircraft (e.g., the thrust level set by the pilot input to engine controller 102) when the icing condition exists.

[0031] Additionally, or alternatively, processor 110 may be configured to instruct other components of turbofan engine 10 to lower the operating line of turbofan engine 10 in response to detection of an icing condition. These additional components may include bleed valves, variable geometry rotors, variable geometry stators, and / or similar structures known in the art for lowering the operating line of a turbofan engine. Typically, as Figure 3 As shown, a normal or nominal position 150 of the operating line of the turbofan engine 10 is set at a specific offset distance from a stall line 152 of the turbofan engine 10 to maximize the efficiency of the turbofan engine 10. Specifically, in some embodiments, the normal or nominal position 150 may be set within a range of about 10% to about 40% below the stall line 152. However, the presence of icing conditions causes a change in the performance of the turbofan engine 10, which moves the operating line from the normal or nominal position 150 to a higher position 154 that is closer to the stall line 152, which may increase the risk of a stall condition by eroding the offset from the stall line 152 and increasing the uncertainty of the position of the operating line relative to the stall line 152. In some embodiments, the higher position 154 may be about 10% to about 20% higher than the normal or nominal position 150. In response to this increase, the processor 110 lowers the higher operating line from 154 to a lower position 156 that is farther away from the stall line 152 than the higher position 154 to offset the increased risk of stalling brought about by the presence of icing conditions. Although in Figure 3 154 as being above the normal or nominal position 150 of the operating line, but lower position 156 may include various positions below upper position 154, including normal or nominal position 150. In some embodiments, lower position 156 may be set to achieve at least 10% additional stall margin compared to upper position 154 of the operating line due to the presence of icing conditions.

[0032] Furthermore, the processor 110 is configured to mitigate some of the effects of icing conditions on the performance of the turbofan engine 10 using electric motor-generators, such as the LP EM 104 and / or the HP EM 106. Specifically, the processor 110 is configured to instruct, via the control interface 116, one or more electric motor-generators to transfer power from a first shaft of the turbofan engine 10 to a second shaft of the turbofan engine 10 in response to the processor 110 determining that an icing condition exists. This power transfer may establish a more conservative operating state for the turbofan engine 10 to anticipate ice shedding, improve thrust response, and prevent stall (e.g., the power transfer lowers the operating line of the turbofan engine 10 to generate more overhead).

[0033] The power transfer between the first shaft and the second shaft is achieved by the processor 110 instructing at least one motor-generator connected to the first shaft gear transmission to apply a negative torque to the first shaft. The negative torque applied by the motor-generator decelerates the first shaft and converts the rotational energy in the first shaft into electrical energy. The electrical energy is then transferred by the motor-generator via an electrical connection to the at least one motor-generator connected to the second shaft gear transmission. After receiving the electrical energy, the motor-generator connected to the second shaft gear transmission converts the electrical energy into a positive torque that is applied to the second shaft. The applied positive torque increases the rotational energy in the second shaft. In some embodiments, the electrical power used to apply the positive torque to the second shaft can be provided in whole or in part by the power supply or storage device 133 (see Figure 2 The power supply or storage device 133 may be supplied by a power source (e.g., a battery) instead of or in addition to the power supply from the motor-generator geared to the first shaft. Various methods are possible for the processor 110 to direct the operation of the motor-generator. Such methods include the processor 110 modulating a voltage or current value applied to a control input of the motor-generator being directed and / or modulating a voltage or current value applied to a bus interconnected with the control input.

[0034] When the icing condition corresponds to an ice accretion event, the first shaft comprises a fan shaft (e.g., Figure 1 LP spool 36) and the second shaft comprises the core shaft of the turbofan engine 10 (e.g. Figure 1HP spool 34). In contrast, when the icing condition corresponds to a de-icing event, the first shaft comprises the core shaft and the second shaft comprises the fan shaft. Power transfer from the core shaft to the fan shaft may be initiated after the stall problem from the previous icing event has been completely resolved. In particular, power transfer from the core shaft to the fan shaft may be initiated in a gradually increasing manner and / or after a time delay from detecting a de-icing event in response to detecting the onset of a core speed overshoot. Furthermore, in some embodiments, when the icing condition corresponds to a de-icing event, some or all of the power transferred from the core shaft may be applied to power outlet device 134 (see Figure 1 ) instead of, or in addition to, the fan shaft to a power outlet device 134. The power outlet device 134 may include a resistive device that converts transferred electrical energy into heat, an aircraft anti-icing system, other similar devices, and / or combinations thereof.

[0035] Figure 4 Graph 200 shows the power transfer between the fan shaft and the core shaft initiated by processor 110 in response to an icing condition. In particular, graph 200 shows the fan speed 201 of fan section 14 and LP spool 36, the core speed 202 of HP spool 34, the low pressure component 105 ( Figure 2 ) of low power requirements 204, high voltage components 107 ( Figure 2 ) and a feedback signal 208 received at feedback interface 114. Time period 209 shown in graph 200 begins at an ice accretion period 210, during which ice begins to accumulate on fan section 14 or other areas of turbofan engine 10. Processor 110 uses feedback signal 208 and / or readings from torque sensors 130, 132 to determine at a detection time 212 that an icing condition exists and corresponds to an ice accretion event.

[0036] After first detection time 212, processor 110 initiates power transfer from the fan shaft to the core shaft during a power transfer period 214. Power transfer mitigates the negative effects of ice accretion and prepares the engine for a potential de-icing event. In particular, power transfer can offset speed mismatches during acceleration caused by flow blockage. Graph 200 illustrates power transfer during the corresponding ramp-up and ramp-down of low power demand 204 and high power demand 206 during power transfer period 214. Following power transfer period 214, turbofan engine 10 may enter a high-frequency oscillation period 216 prior to a de-icing event 218. In some embodiments, high-frequency oscillation period 216 includes the de-icing sequence torque resonance mode described herein. Then, once a de-icing event 218 occurs (either naturally, via the de-icing sequence torque resonance mode, and / or via other de-icing methods known in the art) and the stall risk from ice accretion is resolved, processor 110 initiates power transfer away from the core shaft to prevent or limit core speed overshoot (e.g., N2 overshoot) due to significant ice shedding. As described herein, power can be transferred from the core shaft to the fan shaft by correspondingly ramping down and ramping up the low power demand 204 and the high power demand 206, as shown in graph 200. Thus, excess power in the core shaft that causes core speed overshoot is extracted and can be dumped into the fan shaft or power outlet device 134 as described herein.

[0037] In some embodiments, processor 110 is configured to respond to detection of an icing event at detection time 212 by activating the motor-generator to enter a de-icing sequence torque resonance mode, which injects a disturbance or controlled vibration into the fan shaft to excite specific shaft modes to reduce further ice accretion, shed existing ice, and / or trigger a de-icing event 218. The controlled vibration is specifically configured to trigger a de-icing event 218 in the aircraft engine and, in some embodiments, has a frequency in the range of approximately 10 Hz to approximately 100 Hz. Using the motor-generator in this manner can have minimal to no impact on engine thrust output and be transparent to the aircraft pilot. Furthermore, if de-icing is detected or anticipated, the injection of the disturbance can be timed / coordinated with the delivery of power to the high-voltage component 107 to improve core operability. Furthermore, this process can allow for the elimination of conventional de-icing components of turbofan engine 10. In some embodiments, engine actuators can be used to induce similar disturbances in the engine's core shaft.

[0038] As discussed above, feedback signal 208 is monitored by processor 110 to determine the presence of an icing condition (e.g., detecting ice accretion at detection time 212 and detecting de-icing event 218). In particular, Figure 4The feedback signal 208 shown in FIG includes a high frequency portion (eg, from about 10 Hz to about 500 Hz, or in some embodiments, to about 5000 Hz), which the processor 110 combines with the reference data 124 (see FIG. Figure 2 ) utilizes these high frequency components to identify the presence of icing conditions. In particular, icing conditions may occur on the fan shaft of the turbofan engine 10 (e.g., Figure 1 The oscillations induced on the LP spool 36 correspond to oscillations generated in the electrical signal output by the LP-EM 104, which are generally within the high frequency portion of the electrical signal (e.g., above about 10 Hz, as described above). In some embodiments, the feedback signal 208 may be from the LP-EM 104 (see Figure 2 ), but embodiments using multiple feedback signals from LP-EM 104, HP-EM 106, and / or other electric motor-generators of turbofan engine 10 are also possible. In addition, feedback signal 208 may include the current output from LP-EM 104 and / or HP-EM 106 as an indicator of the torque on low voltage component 105 and high voltage component 107.

[0039] Reference data 124 may include a first data set indicating an icing event and a second data set that is completely or partially different from the first data set and indicates a de-icing event 218. In some embodiments, the first data set includes a first baseline. The first baseline may include a set of values or a range of values for feedback signal 208 under normal operating conditions of turbofan engine 10 (e.g., in the absence of an icing condition). In these embodiments, processor 110 identifies an icing event as a current icing condition when processor 110 identifies a fluctuation in feedback signal 208 within a sampling sequence of the high-frequency portion of feedback signal 208. Specifically, the presence of an icing event is identified when the fluctuation increases in amplitude and / or frequency relative to the first baseline. The first baseline may also include a range or a single expected value for the high-frequency portion of the feedback signal as a function of target operating conditions for the aircraft engine, such as thrust input received from a pilot.

[0040] In some embodiments, the second data set includes a second baseline, a time threshold, and an amplitude threshold. In these embodiments, when processor 110 identifies that the value of the high-frequency portion of the feedback signal in the sampling sequence deviates from the second baseline by an amount greater than the amplitude threshold within a time frame less than the time threshold, processor 110 identifies the icing condition as being a de-icing event 218. In some embodiments, the second baseline comprises a rolling average of the high-frequency portion of feedback signal 208 over time. However, in other embodiments, the second baseline is a set of values or a range of values, such as a set of values or a range of values for feedback signal 208 under normal operating conditions of turbofan engine 10 (e.g., in the absence of an icing condition).

[0041] In some embodiments, the first and second data sets may include a Fast Fourier Transform (FFT) or similar spectral analysis of historical or simulated values of the feedback signal 208 under normal operating conditions of the turbofan engine 10. In these embodiments, the processor 110 may compare the spectral analysis of the first and second data sets with the FFT or similar spectral analysis of the feedback signal 208 during operation of the turbofan engine 10 and identify the presence of an icing condition (e.g., an icing or deicing event) upon detecting an amplitude spike at a frequency between approximately 10-500 Hz. Specifically, such a spike in frequency amplitude is measured relative to the frequency analysis recorded in the first and second data sets and indicates frequencies of the feedback signal 208 that are inconsistent with expected frequencies of shaft modes under normal operation of the turbofan engine 10.

[0042] In some embodiments, the first and second data sets may include one or more signal patterns that each record a previously measured or simulated value or series of values of feedback signal 208 when an icing and / or de-icing event occurs. In these embodiments, processor 110 is configured to determine whether a real-time or sampled value of feedback signal 208 received at feedback interface 114 matches one or more signal patterns of the first and second data sets to determine whether an icing condition exists. Specifically, in some embodiments, the first and second data sets may include an FFT or similar spectral analysis of historical or simulated values of feedback signal 208 when an icing and / or de-icing event occurs, which processor 110 may then compare to the FFT or similar spectral analysis of feedback signal 208 during operation of turbofan engine 10 to identify a possible match.

[0043] Now refer to Figure 5 , showing a method for detecting and responding to an aircraft engine (e.g. Figure 1 Method 400 for determining an icing condition on a turbofan engine 10 of a conventional turbofan engine 10. Method 400 includes receiving, at feedback interface 114, feedback signal 208 from one or more electric motor-generators, as indicated at 410. Next, method 400 includes determining, via processor 110, whether feedback signal 208 corresponds to reference data 124 stored in memory 112, such that feedback signal 208 indicates the presence of an icing condition, as indicated at 420. Determining the correspondence between feedback signal 208 and reference data 124 may include any of the various methods described herein. For example, processor 110 may determine the correspondence using a first data set indicating an icing event, a second data set indicating a de-icing event, a first baseline, a second baseline, a time threshold, and / or an amplitude threshold.

[0044] After determining the correspondence, method 400 includes initiating, via processor 110, an icing condition mitigation action, as indicated at 430, in response to the processor identifying the presence of an icing condition. As described herein, the icing condition mitigation action facilitates terminating the icing condition and / or reducing the increased stall probability of the turbofan engine 10 due to the presence of the icing condition. Each of the various embodiments of the icing condition mitigation action may be initiated as part of method 400. For example, method 400 may include, when processor 110 determines that an icing condition exists, processor 110 instructing one or more components 126 to change the current operating conditions of the turbofan engine 10. As described herein, one or more components 126 may include electric motor-generators (e.g., LP-EM 104 and / or HP-EM 106) and other components that may be used to lower the operating line of the turbofan engine 10 to reduce the increased stall probability due to the icing condition and / or terminate or remove the icing condition by, for example, triggering a de-icing event, where the icing condition includes an ice accretion event.

[0045] Further aspects of the present disclosure are provided by the subject matter of the following clauses:

[0046] An engine controller includes: a feedback interface electrically coupled to one or more electric motors and electric generators deployed in an aircraft engine; a processor electrically coupled to the feedback interface; and a memory electrically coupled to the processor, wherein the memory is configured to store reference data indicating the presence of an icing condition, wherein the processor is configured to monitor feedback signals received from the one or more electric motors and electric generators at the feedback interface and, based on a correspondence between the feedback signals and the reference data, identify whether the feedback signals indicate that the icing condition currently exists in the aircraft engine, and wherein the processor is configured to initiate icing condition mitigation actions in response to identifying the presence of the icing condition, wherein the icing condition mitigation actions facilitate terminating the icing condition and / or reducing a stall probability of the aircraft engine that is increased due to the presence of the icing condition.

[0047] An engine controller according to any preceding clause, wherein the reference data comprises i) a first data set indicative of an icing event for the aircraft engine and ii) a second data set different from the first data set and indicative of a de-icing event for the aircraft engine.

[0048] The engine controller of any preceding clause, wherein the first data set comprises a first baseline, wherein the icing condition identified as existing by the processor comprises the icing event when the processor identifies fluctuations of the feedback signal in a sampling sequence of the feedback signal that increase in amplitude and / or frequency relative to the first baseline, wherein the second data set comprises a second baseline, a time threshold, and an amplitude threshold, and wherein the icing condition identified as existing by the processor comprises the de-icing event when the processor identifies that values of the feedback signal in the sampling sequence deviate from the second baseline by an amount greater than the amplitude threshold within a time frame less than the time threshold.

[0049] An engine controller as described in any preceding clause, wherein the first baseline comprises a first range of expected values of the feedback signal as a function of target operating conditions of the aircraft engine, and wherein the second baseline comprises a rolling average of the feedback signal over time.

[0050] The engine controller of any preceding clause, further comprising a control interface electrically coupled to the processor and configured to electrically couple to one or more components of the aircraft engine, wherein the processor is configured to initiate the icing condition mitigation action by instructing the one or more components, via the control interface, to change a current operating condition of the aircraft engine in response to the processor determining that the icing condition exists.

[0051] The engine controller of any preceding clause, wherein the current operating conditions include a pitch of one or more fans and a speed of one or more cores of the aircraft engine, the processor being configured to change the pitch of the one or more fans and the speed of the one or more cores of the aircraft engine while maintaining a current thrust output of the aircraft in response to the processor determining that the icing condition exists.

[0052] The engine controller of any preceding clause, wherein the one or more components include an additional engine component configurable to, at the direction of the processor, reduce an operating line of the aircraft engine in response to the processor determining that the icing condition exists.

[0053] The engine controller of any preceding clause, wherein the one or more components include the one or more electric motor-generators, and wherein the processor is configured to, in response to the processor determining that the icing condition exists, initiate the icing condition mitigation action by instructing the one or more electric motor-generators, via the control interface, to transfer power from a first shaft of the aircraft engine to a second shaft of the aircraft engine.

[0054] The engine controller of any preceding clause, wherein when the icing condition corresponds to an icing event of the aircraft engine, the first shaft comprises a fan shaft of the aircraft engine and the second shaft comprises a core shaft of the aircraft engine.

[0055] The engine controller of any preceding clause, wherein when the icing condition corresponds to a de-icing event of the aircraft engine, the first shaft comprises a core shaft of the aircraft engine and the second shaft comprises a fan shaft of the aircraft engine.

[0056] The engine controller of any preceding clause, wherein the one or more components include the one or more electric motor-generators, and wherein the processor is configured to, in response to the processor determining that the icing condition exists and that the icing condition corresponds to an icing accretion event for the aircraft engine, initiate the icing condition mitigation action by instructing the one or more electric motor-generators, via the control interface, to apply power from a power source to a core shaft of the aircraft engine.

[0057] The engine controller of any preceding clause, wherein the one or more components include the one or more electric motor-generators, and wherein the processor is configured to, in response to the processor determining that the icing condition exists and that the icing condition corresponds to a de-icing event for the aircraft engine, initiate the icing condition mitigation action by instructing the one or more electric motor-generators, via the control interface, to transfer power from a core shaft of the aircraft engine to a power outlet device.

[0058] The engine controller of any preceding clause, wherein the one or more components include the one or more electric motor-generators, and wherein the processor is configured to, in response to the processor determining that the icing condition exists and that the icing condition corresponds to an icing event for the aircraft engine, initiate the icing condition mitigation action by instructing the one or more electric motor-generators, via the control interface, to induce a controlled vibration on a shaft of the aircraft engine, the controlled vibration being configured to trigger a de-icing event for the aircraft engine.

[0059] An engine icing control system includes: an engine controller; and one or more electric motors and electric generators disposed in an aircraft engine and electrically connected to the engine controller; wherein the engine controller is configured to, in response to the engine controller determining, using a status signal received by the engine controller, that an icing condition exists and that the icing condition corresponds to an icing event for the aircraft engine, instruct the one or more electric motors and electric generators to induce controlled vibrations on a shaft of the aircraft engine, the controlled vibrations being configured to trigger a de-icing event for the aircraft engine.

[0060] An engine icing control system as described in any preceding clause, wherein the shaft of the aircraft engine comprises a fan shaft of the aircraft engine.

[0061] An engine icing control system as described in any preceding clause, wherein the status signal is received from a torque sensor deployed in the aircraft engine.

[0062] An engine icing control system as described in any preceding clause, wherein the status signal is a feedback signal received from one or more electric motor-generators.

[0063] A method for detecting and responding to an icing condition on an aircraft engine, the method comprising: receiving, at a feedback interface of an engine controller, a feedback signal from one or more electric motor-generators deployed in the aircraft engine; determining, via a processor, whether the feedback signal corresponds to reference data stored in a memory such that the feedback signal indicates the presence of an icing condition, the memory being electrically coupled to the processor; and initiating, via the processor, an icing condition mitigation action in response to the processor identifying the presence of the icing condition, wherein the icing condition mitigation action facilitates terminating the icing condition and / or reducing a stall probability of the aircraft engine that is increased due to the presence of the icing condition.

[0064] The method of any preceding clause, wherein initiating the icing mitigation action comprises the processor instructing one or more components of the aircraft engine to lower an operating line of the aircraft engine.

[0065] The method of any preceding clause, wherein the reference data comprises a first data set indicative of an icing event on the aircraft engine and a second data set, different from the first data set and indicative of a de-icing event on the aircraft engine, wherein the first data set comprises a first baseline, wherein when the processor identifies fluctuations of the feedback signal in a sampling sequence of the feedback signal that increase in amplitude and / or frequency relative to the first baseline, the icing condition identified by the processor as existing comprises the icing event, wherein the second data set comprises a second baseline, a time threshold, and an amplitude threshold, and wherein when the processor identifies that values of the feedback signal in the sampling sequence deviate from the second baseline by an amount greater than the amplitude threshold within a time frame less than the time threshold, the icing condition identified by the processor as existing comprises the de-icing event.

[0066] The system and method of any preceding clause, wherein the nominal position of the operating line of the aircraft engine lies within a range of about 10% to about 40% below a stall line of the aircraft engine.

[0067] The system and method of any preceding clause, wherein the presence of the icing condition causes the operating line to be positioned in a range of about 10% to about 20% above the nominal position.

[0068] The system and method of any preceding clause, wherein the condition mitigating action lowers the position of the operating line by a range of approximately 10% compared to an elevated position resulting from the presence of the icing condition.

[0069] The system and method of any preceding clause, wherein the reference data comprises a spectral analysis of historical or simulated values of the feedback signal under normal operating conditions of the aircraft engine.

[0070] The system and method of any preceding clause, wherein the reference data comprises a spectral analysis of historical or simulated values of the feedback signal when the icing condition existed in the aircraft engine.

Claims

1. An engine controller, characterized in that: include: a feedback interface electrically coupled to one or more electric motor-generators deployed in the aircraft engine; and a processor electrically coupled to the feedback interface; a memory electrically coupled to the processor; wherein the memory is configured to store reference data indicating the presence of icing conditions, wherein the processor is configured to monitor feedback signals received at the feedback interface from the one or more electric motor-generators and identify the presence of the icing condition based on a combination of the feedback signals and the reference data; wherein the reference data comprises a first data set indicative of an icing event of the aircraft engine and a second data set different from the first data set and indicative of a deicing event of the aircraft engine; and wherein the first data set includes a first baseline, wherein when the processor identifies that fluctuations of the feedback signal in a sampling sequence of the feedback signal increase in amplitude and / or frequency relative to the first baseline, the icing condition identified as existing by the processor comprises the icing event, wherein the second data set includes a second baseline, a time threshold, and an amplitude threshold, and wherein when the processor identifies that values of the feedback signal in the sampling sequence deviate from the second baseline by an amount greater than the amplitude threshold within a time frame less than the time threshold, the icing condition identified as existing by the processor comprises the de-icing event.

2. The engine controller according to claim 1, characterized in that: Wherein the first baseline comprises a first range of expected values of the feedback signal as a function of currently desired operating conditions of the aircraft engine, and wherein the second baseline comprises a rolling average of the feedback signal over time.

3. An engine controller, characterized in that: include: a feedback interface electrically coupled to one or more electric motor-generators deployed in the aircraft engine; a processor electrically coupled to the feedback interface; a memory electrically coupled to the processor; wherein the memory is configured to store reference data indicating the presence of icing conditions, wherein the processor is configured to monitor feedback signals received at the feedback interface from the one or more electric motor-generators and identify the presence of the icing condition based on a combination of the feedback signals and the reference data; wherein the engine controller further comprises a control interface electrically coupled to the processor and configured to electrically couple to one or more components of the aircraft engine, wherein the processor is configured to instruct the one or more components via the control interface to change a current operating condition of the aircraft engine in response to the processor determining that the icing condition exists; and Wherein the current operating conditions include a pitch of one or more fans and a speed of one or more cores of the aircraft engine, the processor being configured to change the pitch of the one or more fans and the speed of the one or more cores of the aircraft engine while maintaining a current thrust output of the aircraft in response to the processor determining that the icing condition exists.

4. An engine controller, characterized in that: include: a feedback interface electrically coupled to one or more electric motor-generators deployed in the aircraft engine; a processor electrically coupled to the feedback interface; a memory electrically coupled to the processor; wherein the memory is configured to store reference data indicating the presence of icing conditions, wherein the processor is configured to monitor feedback signals received at the feedback interface from the one or more electric motor-generators and identify the presence of the icing condition based on a combination of the feedback signals and the reference data; wherein the engine controller further comprises a control interface electrically coupled to the processor and configured to electrically couple to one or more components of the aircraft engine, wherein the processor is configured to instruct the one or more components via the control interface to change a current operating condition of the aircraft engine in response to the processor determining that the icing condition exists; and Wherein the one or more components include an additional engine component, the additional engine component can be configured to, in response to the processor determining that the icing condition exists, reduce an operating line of the aircraft engine at the direction of the processor.

5. An engine controller, characterized in that: include: a feedback interface electrically coupled to one or more electric motor-generators deployed in the aircraft engine; a processor electrically coupled to the feedback interface; a memory electrically coupled to the processor; wherein the memory is configured to store reference data indicating the presence of icing conditions, wherein the processor is configured to monitor feedback signals received at the feedback interface from the one or more electric motor-generators and identify the presence of the icing condition based on a combination of the feedback signals and the reference data; wherein the engine controller further comprises a control interface electrically coupled to the processor and configured to electrically couple to one or more components of the aircraft engine, wherein the processor is configured to instruct the one or more components via the control interface to change a current operating condition of the aircraft engine in response to the processor determining that the icing condition exists; and wherein the one or more components include the one or more motors and motor-generators, and wherein the processor is configured to initiate the icing condition mitigation action by instructing the one or more motors and motor-generators via the control interface to perform an operation, the operation being one of the following: responsive to the processor determining that the icing condition exists, transferring power from a first shaft of the aircraft engine to a second shaft of the aircraft engine; applying power from a power source to a core shaft of the aircraft engine in response to the processor determining that the icing condition exists and that the icing condition corresponds to an icing event of the aircraft engine; in response to the processor determining that the icing condition exists and that the icing condition corresponds to a de-icing event for the aircraft engine, transferring power from a core shaft of the aircraft engine to a power outlet device; or In response to the processor determining that the icing condition exists and that the icing condition corresponds to an icing event for the aircraft engine, a controlled vibration is induced on a shaft of the aircraft engine, the controlled vibration configured to trigger a de-icing event for the aircraft engine.

6. The engine controller according to claim 5, characterized in that: Wherein when the icing condition corresponds to an ice accretion event of the aircraft engine, the first shaft comprises a fan shaft of the aircraft engine and the second shaft comprises a core shaft of the aircraft engine.

7. The engine controller according to claim 5, characterized in that: Wherein when the icing condition corresponds to a de-icing event of the aircraft engine, the first shaft comprises a core shaft of the aircraft engine and the second shaft comprises a fan shaft of the aircraft engine.

8. An engine icing control system, characterized in that: include: engine controller; and one or more electric motor-generators disposed in an aircraft engine and electrically connected to the engine controller; wherein the engine controller is configured to, in response to the engine controller determining, using status signals received by the engine controller, that an icing condition exists and that the icing condition corresponds to an icing event for the aircraft engine, instruct the one or more electric motor-generators to induce a controlled vibration on a shaft of the aircraft engine, the controlled vibration being configured to trigger a de-icing event for the aircraft engine.

9. The engine icing control system according to claim 8, characterized in that: Wherein the shaft of the aircraft engine comprises a fan shaft of the aircraft engine.

10. The engine icing control system according to claim 8, characterized in that: The status signal is received from a torque sensor deployed in the aircraft engine.

11. The engine icing control system according to claim 8, characterized in that: The status signal is a feedback signal received from one or more motor generators.

12. A method for detecting and responding to icing conditions on an aircraft engine, characterized in that: The method comprises: receiving, at a feedback interface of the engine controller, feedback signals from one or more electric motor-generators deployed in an aircraft engine; determining, via a processor, whether the feedback signal corresponds to reference data stored in a memory electrically coupled to the processor such that the feedback signal indicates an icing condition exists; and initiating, via the processor, an icing condition mitigation action in response to the processor identifying the presence of the icing condition, wherein the icing condition mitigation action facilitates terminating the icing condition and / or reducing a stall probability of the aircraft engine that is increased due to the presence of the icing condition; wherein the reference data includes a first data set indicating an icing event of the aircraft engine and a second data set different from the first data set and indicating a de-icing event of the aircraft engine; wherein the first data set includes a first baseline, wherein when the processor identifies that fluctuations of the feedback signal in a sampling sequence of the feedback signal increase in amplitude and / or frequency relative to the first baseline, the icing condition identified by the processor as existing includes the icing event; wherein the second data set includes a second baseline, a time threshold, and an amplitude threshold, and wherein when the processor identifies that values of the feedback signal in the sampling sequence deviate from the second baseline by an amount greater than the amplitude threshold within a time frame less than the time threshold, the icing condition identified by the processor as existing includes the de-icing event.

13. The method according to claim 12, characterized in that Wherein initiating the icing mitigation action includes the processor instructing one or more components of the aircraft engine to lower an operating line of the aircraft engine.

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