Patent ID: 12258139

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosed apparatus and method are presented herein by way of exemplification and not limitation with reference to the Figures.

FIG.1schematically illustrates a gas turbine engine20. The gas turbine engine20is disclosed herein as a two-spool turbofan that generally incorporates a fan section22, a compressor section24, a combustor section26and a turbine section28. The fan section22drives air along a bypass flow path B in a bypass duct, while the compressor section24drives air along a core flow path C for compression and communication into the combustor section26then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including single-spool and three-spool architectures.

The exemplary engine20generally includes a low speed spool30and a high speed spool32mounted for rotation about an engine central longitudinal axis A relative to an engine static structure36via several bearing systems38. It should be understood that various bearing systems38at various locations may alternatively or additionally be provided, and the location of bearing systems38may be varied as appropriate to the application.

The low speed spool30generally includes an inner shaft40that interconnects a fan42, a low pressure compressor44and a low pressure turbine46. The inner shaft40is connected to the fan42through a speed change mechanism, which in exemplary gas turbine engine20is illustrated as a geared architecture48to drive the fan42at a lower speed than the low speed spool30. The high speed spool32includes an outer shaft50that interconnects a high pressure compressor52and high pressure turbine54. A combustor56is arranged in exemplary gas turbine engine20between the high pressure compressor52and the high pressure turbine54. An engine static structure36is arranged generally between the high pressure turbine54and the low pressure turbine46. The engine static structure36further supports bearing systems38in the turbine section28. The inner shaft40and the outer shaft50are concentric and rotate via bearing systems38about the engine central longitudinal axis A which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor44then the high pressure compressor52, mixed and burned with fuel in the combustor56, then expanded over the high pressure turbine54and low pressure turbine46. The turbines46,54rotationally drive the respective low speed spool30and high speed spool32in response to the expansion. It will be appreciated that each of the positions of the fan section22, compressor section24, combustor section26, turbine section28, and fan drive gear system48may be varied. For example, gear system48may be located aft of combustor section26or even aft of turbine section28, and fan section22may be positioned forward or aft of the location of gear system48.

The engine20in one example is a high-bypass geared aircraft engine. In a further example, the engine20bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture48is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine46has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine20bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor44, and the low pressure turbine46has a pressure ratio that is greater than about five 5:1. Low pressure turbine46pressure ratio is pressure measured prior to inlet of low pressure turbine46as related to the pressure at the outlet of the low pressure turbine46prior to an exhaust nozzle. The geared architecture48may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section22of the engine20is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and 35,000 ft (10,688 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)]{circumflex over ( )}0.5. The “Low corrected fan tip speed” as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

While the example ofFIG.1illustrates one example of the gas turbine engine20, it will be understood that any number of spools, inclusion or omission of the gear system48, and/or other elements and subsystems are contemplated. Further, rotor systems described herein can be used in a variety of applications and need not be limited to gas turbine engines for aircraft applications.

Referring now toFIG.2, a schematic illustration of an aircraft100includes an engine system200with first and second gas turbine engines20A,20B as embodiments of the gas turbine engine20ofFIG.1. Although the engine system200is depicted with two gas turbine engines20, it will be understood that the engine system200can include additional gas turbine engines (e.g., three or more instances of the gas turbine engine20on the aircraft100). Continuing with a two-engine example for purposes of explanation, each of the first and second gas turbine engines20A,20B can have an associated engine controller204A,204B. The engine controllers204A,204B can receive commands and data from an aircraft controller206of the aircraft100. The aircraft controller206can accept input from and provide output to a pilot interface205. The pilot interface205can include redundant instances of switches, knobs, levers, buttons, displays, and the like, which can be controlled by a pilot and/or co-pilot of the aircraft100. Collectively, the engine controllers204A,204B and aircraft controller206may be referred to as a control system202. Control logic and command generation can be implemented by any portion of the control system202and may be distributed, for example, between the engine controllers204A,204B and/or aircraft controller206. The aircraft controller206may receive pilot inputs though pilot interface205and control multiple aspects of the aircraft100. Examples of pilot inputs can be normal operating commands and/or override commands. For instance, a pilot can override a motoring sequence, as described herein, based on various conditions, such as, a flight delay, a return-to-gate condition, a maintenance condition, an engine shutdown condition, and other such factors. The aircraft controller206and/or engine controllers204A,204B can also support automated modes of operation, such as an auto-start mode, an auto-cooldown mode, an auto-pilot mode, and other such modes. The engine controllers204A,204B can be embodied in one or more full authority digital engine controls (FADECs), for example.

The engine controllers204A,204B and aircraft controller206can each include memory to store instructions that are executed by one or more processors on one or more channels. The executable instructions may be stored or organized in any manner and at any level of abstraction, such as in connection with a controlling and/or monitoring operation of the first and second gas turbine engines20A,20B. The one or more processors can be any type of central processing unit (CPU), including a general purpose processor, a digital signal processor (DSP), a microcontroller, an application specific integrated circuit (ASIC), a field programmable gate array (FPGA), or the like. Also, in embodiments, the memory may include random access memory (RAM), read only memory (ROM), or other electronic, optical, magnetic, or any other computer readable medium onto which is stored data and control algorithms in a non-transitory form.

In the example ofFIG.2, the engine controllers204A,204B can meter a fuel flow to the combustor section26(FIG.1) of respective gas turbine engines20A,20B based on a current operating mode of the aircraft100according to one or more fuel burn schedules. For example, the engine controllers204A,204B can meter a fuel flow to the combustor section26according to an idle fuel burn schedule, a take-off fuel burn schedule, a climb fuel burn schedule, a cruise fuel burn schedule, a descent fuel burn schedule, and a landing fuel burn schedule. The characteristics of the fuel burn schedules may be adjusted for aircraft and environmental conditions. To reduce fuel consumption while operating on the ground, embodiments perform taxiing of the aircraft100absent fuel burn by the gas turbine engines20A,20B. Each of the gas turbine engines20A,20B can have a starter216A,216B that is configured to drive rotation of the high speed spool32within each of the gas turbine engines20A,20B respectively. Driving rotation of the high speed spool32can increase pressure in the combustor56to support ignition by ignitors and subsequent fuel burn.

In embodiments, the starters216A,216B can be driven in a dry cranking mode, where fuel burn is inhibited, to establish an airflow through the gas turbine engines20A,20B. Although the high speed spool32is not directly coupled to the fan42, by motoring the high speed spool32at a motoring speed above an idle speed of the gas turbine engines20A,20B, the airflow drawn into the gas turbine engines20A,20B can also result in driving rotation of the low speed spool30and the fan42to provide propulsion during the taxi mode. Idle speed can refer to a lowest speed at which the gas turbine engines20A,20B typically operate with fuel burn when on the ground with a minimum fuel burn thrust. In some embodiments, a motor215A,215B can be coupled to the low speed spool30to provide a supplemental or alternate source of rotational power to the low speed spool30. For instance, the sizing of the motors215A,215B can be reduced when configured for use in combination with the starters216A,216B as compared to relying only upon the motors215A,215B for taxi mode operations. Depending upon the component sizing and available power, the starters216A,216B and/or motors215A,215B can be operated sequentially or in parallel. For instance, where different power sources are used, the aircraft100may taxi with gas turbine engine20A being driven by starter216A, for instance through pneumatic power, and with gas turbine engine20B being driven by motor215B, for instance through electric power. The use of power sources between engine pairs can alternate to balance component life.

Various power source options for the starters216A,216B and/or motors215A,215B can include a power source208and/or a stored power source214A,214B. In the example ofFIG.2, the power source208is depicted as providing input power218A,218B to the starters216A,216B respectively. For instance, the power source208can be an auxiliary power unit, an electric generator, a hydraulic source, a pneumatic source, and/or another source of power known in the art. The stored power sources214A,214B can be a battery system, supercapacitors, ultracapacitors, a flywheel system, or other such forms of stored power. The stored power sources214A,214B can be recharged, for example, by power supplied from the power source208. In some embodiments, the motors215A,215B can be motor/generators, where the motors215A,215B may be operated in a generator mode to store power in the stored power sources214A,214B. Further, the starters216A,216B can be motor/generators, which may receive power from or deliver power to the stored power sources214A,214B. The motors215A,215B and starters216A,216B can be implemented using a variety of sources of power. For instance, the motors215A,215B and/or starters216A,216B can include one or more of electric motor, a pneumatic drive, and a hydraulic drive powered by one or more of: an auxiliary power unit (APU), a battery system, an electric generator, a hydraulic source, and a pneumatic source. Further, although depicted in close proximity to the gas turbine engines20A,20B, the motors215A,215B and/or starters216A,216B can be distributed at other locations of the aircraft100. For instance, where the power source208is an APU, a motor/generator can be driven by the APU to provide taxi-mode and/or starting power for the gas turbine engines20A,20B.

Where the speed of the high speed spool32is increased above an idle speed to draw in airflow during taxiing prior to fuel combustion, one or more engine bleeds210A,210B of the gas turbine engines20A,20B can be controlled to extract an engine bleed flow212A,212B to reduce pressure within the gas turbine engines20A,20B prior to starting. Other approaches to reducing pressure within the gas turbine engines20A,20B before starting can include reducing/stopping input power218A,218B by controlling one or more valves, venting pneumatic power, and/or controlling the power source208to reduce/prevent delivery of the input power218A,218B.

During normal operation, heating within the first and second gas turbine engines20A,20B can result in thermal distortion of one or more spools (e.g., low speed spool30and/or high speed spool32ofFIG.1) during engine shutdown such that on restart, vibration and/or rubbing of blade tips within an engine casing can result, particularly where the rotational speed increases towards a major resonance speed (which may be referred to as a critical speed). Performing taxiing operations by driving rotation within the gas turbine engines20A,20B without fuel burn active can result in reducing internal temperatures of the gas turbine engines20A,20B to reduce/prevent bowed rotor conditions and avoid rubbing as the rotational speed increases up to and beyond the critical speed.

The starters216A,216B and/or motors215A,215B can interface to respective engines20A,20B through gear trains, gearboxes, shafts, clutches, and/or other interfaces that provide the starters216A,216B and/or motors215A,215B with a linkage to control rotation of an engine core of the first and second gas turbine engines20A,20B. For instance, the starters216A,216B and/or motors215A,215B can each be an electric motor that results in rotation of the first and second gas turbine engines20A,20B at targeted speeds to support taxiing of the aircraft100, as one example. Alternatively, the starters216A,216B can be pneumatic starters, such as an air turbine starter, that rotates components of the first and second gas turbine engines20A,20B at targeted speeds. As a further alternative, the starters216A,216B and/or motors215A,215B can be driven to rotate by pressurized hydraulic fluid.

A speed of the low speed spool30(also referred to as N1 speed) can be monitored by engine controllers204A,204B and/or through the aircraft controller206and/or through the pilot interface205, where the speed of the low speed spool30serves as a proxy for taxi speed. Further or alternatively, other sensed or derived values can be used to determine a taxi speed of the aircraft100, such as wheel speed. The speed of the high speed spool32can be increased or decreased based on the speed of the low speed spool30and/or based on a taxi speed of the aircraft100. The starters216A,216B can be controlled in response to the speed of the low speed spool30and/or based on an observed taxi speed of the aircraft100. Driving rotation of the high speed spool32may not map precisely to a taxi speed, since the fan42ofFIG.1provides the majority of ground-based propulsion for taxi operations, and the high speed spool32is not directly coupled to the fan42. When the low speed spool30is not directly driven by the motors215A,215B, the rotation of the high speed spool32driven by the starters216A,216B can result in rotation of the low speed spool30as air is drawn into the engine core in response to rotation of the high speed spool32, thereby resulting in rotation of the fan42. Where the starters216A,216B are pneumatic or hydraulic, the rotational speed may be controlled by one or more valves. For instance, a speed reduction may be achieved by opening one or more bleed valves, partially closing one or more flow control valves, applying pulse-width modulation to open and close valves with a targeted duty cycle, and/or reducing a source of input fluid, such as air from an auxiliary power unit. Speed increases provided by the starters216A,216B can be performed by closing one or more bleed valves, opening one or more flow control valves, modifying a duty cycle where pulse-width modulation is used for valve control, and/or increasing a source of input fluid, such as air from an auxiliary power unit. Where the starters216A,216B are electrically driven, the speed can be controlled by increasing or decreasing electrical current provided to the starters216A,216B. Using the speed of the low speed spool30and/or the taxi speed of the aircraft100as a control input for the starters216A,216B that drive rotation of the high speed spool32can provide a wider range of flexibility in controlling the power delivery to the starters216A,216B as the rotational speed of the high speed spool32(also referred to as N2 speed) need not be precisely controlled to a specific target value.

FIG.3depicts a sequence diagram300for operation of the engine system200of aircraft100with further reference toFIGS.1-2. Time progresses from left to right onFIG.3and is not to scale; rather, the sequence diagram300is a sequential illustration for purposes of explanation.

In the example ofFIG.3, a pushback event304of the aircraft100can be performed, for instance, while the engines20A,20B are depowered. During time305, the aircraft100may be moved by an external force, such as a ground-based tug or cart. Alternatively, the pushback event304can include moving a gate or stairs while the aircraft100remains stationary. At power-on event306, one or more actuation systems can be powered on. For instance, the power source208(e.g., an APU) can power one or more electric actuators to provide electric power to the starters216A,216B and/or motors215A,215B. Alternatively, the power-on event306can provide pneumatic or hydraulic power for the starters216A,216B and/or motors215A,215B. During time307, power can be conditioned until power delivery is ready at event308, for instance, by accelerating electric generators, setting contactor positions, and the like. At event308, power can be delivered to the starters216A,216B and/or motors215A,215B. During time309, power delivery to the starters216A,216B and/or motors215A,215B can be at an amount to establish rotation of the engine core within either or both of the gas turbine engines20A,20B but may not be at a high enough level to taxi the aircraft100. At event310, a taxi mode is initiated to reposition the aircraft100on the ground. During time311, either or both of the gas turbine engines20A,20B are motored, absent fuel burn, using the starters216A,216B and/or motors215A,215B. A motoring speed of either or both of the gas turbine engines20A,20B can be accelerated, absent fuel burn, above an idle speed to provide propulsion during the taxi mode. At event312, a targeted new position of the aircraft100is reached. During time313, the motoring speed of either or both of the gas turbine engines20A,20B can be decreased, absent fuel burn. For instance, the starters216A,216B and/or motors215A,215B may continue to rotate components of the gas turbine engines20A,20B to drive accessories and keep the aircraft100is a ready state. At event314, a shutdown process may be initiated where the actuators or power provided to the starters216A,216B and/or motors215A,215B is reduced to an off state. During time315, depowering of the power source208can occur. At event316, the depowering of the power source208can be completed. At time317, rotation within the gas turbine engines20A,20B is reduced until rotation ceases at event318and engine shutdown completes.

FIG.4depicts a sequence diagram400for operation of the engine system200of aircraft100with further reference toFIGS.1-2. Time progresses from left to right onFIG.4and is not to scale; rather, the sequence diagram400is a sequential illustration for purposes of explanation. The example ofFIG.4is described with respect to using the starters216A,216B, for instance, in embodiments where the motors215A,215B are not installed or used.

In the example ofFIG.4, a pushback event404of the aircraft100can be performed, for instance, while the engines20A,20B are depowered. During time405, the aircraft100may be moved by an external force, such as a ground-based tug or cart. Alternatively, the pushback event404can include moving a gate or stairs while the aircraft100remains stationary. At an initiate taxi event406, one or more actuation systems can be powered on. For instance, the power source208(e.g., an APU) can power one or more electric actuators to provide electric or pneumatic power to the starters216A,216B. During time407, the starters216A,216B can operate absent fuel burn to motor either or both of the gas turbine engines20A,20B during taxi mode. At event408, the pilot or co-pilot may decide to activate an auto-start mode of either or both of the gas turbine engines20A,20B during taxiing, for instance through pilot interface205. The engine controllers204A,204B can initiate an engine start sequence of either or both of the gas turbine engines20A,20B based on detecting an auto-start mode selection as a change in starting mode (e.g., from a previous selection of manual starting control). The engine controllers204A,204B can control the starters216A,216B, fuel flow, and ignitors of either or both of the gas turbine engines20A,20B to start fuel combustion during the taxi mode based on detecting the auto-start mode selection. The engine controllers204A,204B can also control the engine bleeds210A,210B and/or other aspects of the gas turbine engines20A,20B selected for automated start sequencing. Upon reaching necessary conditions for starting, either or both of the gas turbine engines20A,20B can switch to a fuel-burning mode of operation during taxi mode in time409. At event410, the aircraft100can reach a targeted new position and may proceed during time411to resume taxiing, transition to take-off, or perform a shutdown.

FIG.5depicts a sequence diagram500for operation of the engine system200of aircraft100with further reference toFIGS.1-2. Time progresses from left to right onFIG.5and is not to scale; rather, the sequence diagram500is a sequential illustration for purposes of explanation. The example ofFIG.5is described with respect to using both the motors215A,215B and the starters216A,216B, for instance, in embodiments where the motors215A,215B are sized to provide power to drive rotation of fan42during taxi mode.

In the example ofFIG.5, a pushback event504of the aircraft100can be performed, for instance, while the engines20A,20B are depowered. During time505, the aircraft100may be moved by an external force, such as a ground-based tug or cart. Alternatively, the pushback event504can include moving a gate or stairs while the aircraft100remains stationary. At an initiate taxi event506, one or more motors215A,215B can be powered by the power source208and/or stored power source214A,214B. During time507, the motors215A,215B can operate absent fuel burn to motor either or both of the gas turbine engines20A,20B during taxi mode. At event508, the pilot or co-pilot may decide to activate an auto-start mode of either or both of the gas turbine engines20A,20B during taxiing, for instance through pilot interface205. The engine controllers204A,204B can initiate an engine start sequence of either or both of the gas turbine engines20A,20B based on detecting an auto-start mode selection as a change in starting mode (e.g., from a previous selection of manual starting control). The engine controllers204A,204B can control the starters216A,216B, fuel flow, and ignitors of either or both of the gas turbine engines20A,20B to start fuel combustion during the taxi mode based on detecting the auto-start mode selection. The engine controllers204A,204B can also control the engine bleeds210A,210B and/or other aspects of the gas turbine engines20A,20B selected for automated start sequencing. In some embodiments, the motors215A,215B can remain active to continue motoring of the fan42during the starting process in time509. As fuel burn commences, the output of the motors215A,215B can be tapered off. At event510, the aircraft100can reach a targeted new position and may proceed during time511to resume taxiing, transition to take-off, or perform a shutdown.

FIG.6depicts a sequence diagram600for operation of the engine system200of aircraft100with further reference toFIGS.1-2. Time progresses from left to right onFIG.6and is not to scale; rather, the sequence diagram600is a sequential illustration for purposes of explanation. The example ofFIG.6is described with respect to using the motors215A,215B, for instance, in embodiments where the starters216A,216B are not installed or used.

In the example ofFIG.6, a pushback event604of the aircraft100can be performed, for instance, while the engines20A,20B are depowered. During time605, the aircraft100may be moved by an external force, such as a ground-based tug or cart. Alternatively, the pushback event604can include moving a gate or stairs while the aircraft100remains stationary. At an initiate taxi event606, one or more motors215A,215B can be powered by the power source208and/or stored power source214A,214B. During time607, the motors215A,215B can operate absent fuel burn to motor either or both of the gas turbine engines20A,20B during taxi mode. At event608, the pilot or co-pilot may decide to activate an auto-start mode of either or both of the gas turbine engines20A,20B during taxiing, for instance through pilot interface205. The engine controllers204A,204B can initiate an engine start sequence of either or both of the gas turbine engines20A,20B based on detecting an auto-start mode selection as a change in starting mode (e.g., from a previous selection of manual starting control). The engine controllers204A,204B can control the fuel flow, and ignitors of either or both of the gas turbine engines20A,20B to start fuel combustion during the taxi mode based on detecting the auto-start mode selection. The engine controllers204A,204B can also control the engine bleeds210A,210B and/or other aspects of the gas turbine engines20A,20B selected for automated start sequencing. Upon reaching necessary conditions for starting, either or both of the gas turbine engines20A,20B can switch to a fuel-burning mode of operation during taxi mode in time609. As fuel burn commences, the output of the motors215A,215B can be tapered off. At event610, the aircraft100can reach a targeted new position and may proceed during time611to resume taxiing, transition to take-off, or perform a shutdown.

Referring now toFIG.7with continued reference toFIGS.1-6,FIG.7is a flow chart illustrating a method700of engine system control in accordance with an embodiment. The method700may be performed, for example, by the engine system200ofFIG.2. For purposes of explanation, the method700is described primarily with respect to the engine system200ofFIG.2; however, it will be understood that the method700can be performed on other configurations (not depicted).

At block702, the control system202can motor the gas turbine engine20A,20B, absent fuel burn, during a taxi mode of the aircraft100. Motoring of the gas turbine engine20A,20B can include the use of one or more of the starters216A,216B and/or motors215A,215B to drive the fan42of engines20A,20B to rotate directly or indirectly. At block704, the control system202can accelerate a motoring speed of the gas turbine engine20A,20B, absent fuel burn, above an idle speed of the gas turbine engine20A,20B to provide propulsion during the taxi mode. The motoring speed can be driven by one or more of: an electric motor, a pneumatic drive, and a hydraulic drive powered by one or more of: an auxiliary power unit, a battery system, an electric generator, a hydraulic source, and a pneumatic source. For instance, the motor215A,215B used for motoring can be a motor/generator that receives power from and charges stored power source214A,214B. Further, the starter216A,216B can be electrically driven, pneumatically driven, or hydraulicly driven by the power source208for motoring. The motoring speed can be reached based on controlling the starter216A,216B to dry motor the gas turbine engine20A,20B and/or controlling an electric motor (e.g., motor215A,215B) of the gas turbine engine20A,20B. In some aspects, the starters216A,216B drive rotation of the high speed spool32of engines20A,20B based on a speed of the low speed spool30of engines20A,20B and/or based on an observed taxi speed of the aircraft100. Taxi operations can be performed using only the starters216A,216B driving rotation of the high speed spool32of engines20A,20B to induce rotation of the fan42absent a physical coupling between the high speed spool32and the fan42of engines20A,20B.

At block706, the control system202can decrease the motoring speed of the gas turbine engine20A,20B, absent fuel burn, based on a change in a starting mode of the gas turbine engine20A,20B or the aircraft100reaching a targeted new position. The control system202can be configured to power one or more electric actuators to drive the motoring speed and depower the one or more electric actuators after reaching a targeted new position. The control system202can be further configured to decrease the motoring speed of the gas turbine engine20A,20B to the idle speed or a below idle speed after reaching the targeted new position. The reaching of a targeted new position can be determined based on input received through the pilot interface205and/or instrumentation of the aircraft100. The decrease of the motoring speed of the gas turbine engine20A,20B can be performed by a reduction in pressure delivered to a starter216A,216B of the gas turbine engine20A,20B.

In embodiments, the control system202can initiate an engine start sequence of the gas turbine engine20A,20B based on detecting an auto-start mode selection as the change in the starting mode. The control system202can control a starter216A,216B, a motor215A,215B, fuel flow, and ignitors of the gas turbine engine20A,20B to start fuel combustion during the taxi mode based on detecting the auto-start mode selection.

In some embodiments, the gas turbine engine20A and gas turbine engine20B can be separately controlled. For example, starter216A can be used to provide a first motoring speed of the gas turbine engine20A, and motor215B can be used to provide a second motoring speed of the gas turbine engine20B during the taxi mode. The sequence and selection of motoring sources can be alternated between taxiing events. Additionally, the control system202can operate the engine system200when some components, such as the engine controllers204A,204B are depowered. For instance, pilot commands issued through the pilot interface205can trigger the aircraft controller206to control aspects of the motor215A,215B, starter216A,216B, and/or power source208when either or both of the engine controllers204A,204B are depowered. Embodiments can include control paths operable through the engine controllers204A,204B and/or through other inputs, such as pilot overrides that allow one or more aspects of the motor215A,215B, starter216A,216B, and/or power source208to be controlled separate from the engine controllers204A,204B. Control actions can be performed sequentially or in parallel per engine or engine group.

Also, while the above description describes a process for a twin-engine aircraft, a similar procedure can be applied to aircraft with more than two engines. For example, in the case of more than two engines, more than one engine (e.g., one or more additional gas turbine engine20) may be motored during taxiing.

While the above description has described the flow process ofFIG.7in a particular order, it should be appreciated that unless otherwise specifically required in the attached claims that the ordering of the steps may be varied. Further, the designation of the first and second gas turbine engines20A,20B can be arbitrary and need not map to a particular engine (e.g., left or right) as the designations can change such that at least one of the gas turbine engines20A,20B is motored during portions of taxi operations.

The term “about” is intended to include the degree of error associated with measurement of the particular quantity based upon the equipment available at the time of filing the application.

The terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the present disclosure. As used herein, the singular forms “a”, “an” and “the” are intended to include the plural forms as well, unless the context clearly indicates otherwise. It will be further understood that the terms “comprises” and/or “comprising,” when used in this specification, specify the presence of stated features, integers, steps, operations, elements, and/or components, but do not preclude the presence or addition of one or more other features, integers, steps, operations, element components, and/or groups thereof.

While the present disclosure has been described with reference to an exemplary embodiment or embodiments, it will be understood by those skilled in the art that various changes may be made and equivalents may be substituted for elements thereof without departing from the scope of the present disclosure. In addition, many modifications may be made to adapt a particular situation or material to the teachings of the present disclosure without departing from the essential scope thereof. Therefore, it is intended that the present disclosure not be limited to the particular embodiment disclosed as the best mode contemplated for carrying out this present disclosure, but that the present disclosure will include all embodiments falling within the scope of the claims.