Patent Publication Number: US-2023137815-A1

Title: Hybrid electric single engine descent mode activation logic

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims the benefit of U.S. Provisional Patent Application No. 63/273,471 filed Oct. 29, 2021, the disclosure of which is incorporated herein by reference in its entirety. 
    
    
     BACKGROUND 
     The subject matter disclosed herein generally relates to rotating machinery and, more particularly, to a method and an apparatus for a hybrid electric single engine descent mode activation logic. 
     Gas turbine engines are typically inefficient to operate at low power settings. Operation of a gas turbine engine at idle is the typical lowest power setting available once the gas turbine engine has been started. In some instances, thrust produced at idle may be greater than the thrust needed for ground-based operations, such as taxiing and waiting in a parked position prior to takeoff or after landing. This can result in excess fuel consumption and may reduce engine component life with many repeated taxi, takeoff, and landing cycles. 
     In a hybrid gas turbine engine, an electric motor can be available to assist the gas turbine engine operation by adding rotational force to a spool of the gas turbine engine while fuel flow to the gas turbine engine is reduced below idle or shut off. Such a configuration can result in non-intuitive control from a pilot perspective, depending on how the two energy sources, fuel and electricity, are expected to be managed through the range of aircraft operation. In some control configurations, during operations such as engine start, thrust control may not be available to the pilot. 
     BRIEF DESCRIPTION 
     In another exemplary embodiment a system includes a first gas turbine engine of an aircraft, the first gas turbine engine having a first low speed spool, a first high speed spool, and a first combustor. The system further includes a first high spool motor configured to augment rotational power of the first high speed spool. The system further includes a second gas turbine engine of an aircraft, the second gas turbine engine having a second low speed spool, a second high speed spool, and a second combustor. The system further includes a second high spool motor configured to augment rotational power of the second high speed spool. The system further includes a controller. The controller determines a thrust requirement to satisfy the desired glide slope. The controller further determines whether thrust matching can be maintained while operating a first gas turbine engine in a fuel-burning mode and operating a second gas turbine engine in an electrically powered mode. Responsive to determining that thrust matching cannot be maintained, the controller commands fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that the first gas turbine engine provides a first thrust and that the second gas turbine engine provides a second thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that determining whether thrust matching can be maintained includes comparing the first thrust to the second thrust 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that it is determined that thrust matching can be maintained when the second thrust satisfies a threshold difference relative to the first thrust 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that it is determined that thrust matching cannot be maintained when the second thrust fails to satisfy a threshold difference relative to the first thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that the controller is further configured to: determine a glide range to an alternate landing location; and responsive to determining that the glide range exceeds a distance threshold, command fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that the controller is further configured to track a first amount of time the second engine spends in the fuel-burning mode and a second amount of time the second engine spends in the electrically powered mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that commanding the fuel flow to the combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode is based at least in part on at least one of the first amount of time or the second amount of time. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that, responsive to determining that thrust matching can be maintained, the first gas turbine engine continues to operate in the fuel-burning mode and the second gas turbine engine continues to operate in the electrically powered mode 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the system may include that the controller is further configured to receive the desired glide slope. 
     In one exemplary embodiment, a computer-implemented method for a a hybrid electric single engine descent mode activation logic is provided. The method includes determining a thrust requirement to satisfy a desired glide slope. The method further includes determining, based on the thrust requirement, whether thrust matching can be maintained while operating a first gas turbine engine in a fuel-burning mode and operating a second gas turbine engine in an electrically powered mode. The method further includes, responsive to determining that thrust matching cannot be maintained, commanding fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include that the first gas turbine engine provides a first thrust and that the second gas turbine engine provides a second thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include that determining whether thrust matching can be maintained includes comparing the first thrust to the second thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include that it is determined that thrust matching can be maintained when the second thrust satisfies a threshold difference relative to the first thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include that it is determined that thrust matching cannot be maintained when the second thrust fails to satisfy a threshold difference relative to the first thrust. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include determining a glide range to an alternate landing location; and responsive to determining that the glide range exceeds a distance threshold, commanding fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include tracking a first amount of time the second engine spends in the fuel-burning mode and a second amount of time the second engine spends in the electrically powered mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include that commanding the fuel flow to the combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode is based at least in part on at least one of the first amount of time or the second amount of time. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include responsive to determining that thrust matching can be maintained, continuing to operate the first gas turbine engine in the fuel-burning mode and continuing to operate the second gas turbine engine in the electrically powered mode. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the method may include receiving the desired glide slope. 
     The above features and advantages, and other features and advantages, of the disclosure are readily apparent from the following detailed description when taken in connection with the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following descriptions should not be considered limiting in any way. With reference to the accompanying drawings, like elements are numbered alike: 
         FIG.  1    is a schematic diagram of an aircraft including dual hybrid electric propulsions systems, in accordance with an embodiment of the disclosure; 
         FIG.  2    is a schematic diagram of a hybrid electric propulsion system, in accordance with an embodiment of the disclosure; 
         FIG.  3    is a schematic diagram of control signal paths of a hybrid electric propulsion system, in accordance with an embodiment of the disclosure; 
         FIG.  4    is a plot that graphically illustrates a relationship between engine spool speeds and time when transitioning through multiple operating modes, in accordance with an embodiment of the disclosure; 
         FIG.  5    is a plot that graphically illustrates a relationship between thrust and throttle lever angle, in accordance with an embodiment of the disclosure; 
         FIG.  6    is a flow chart illustrating a method, in accordance with an embodiment of the disclosure; and 
         FIG.  7    is a flow chart illustrating a method, in accordance with an embodiment of the disclosure. 
     
    
    
     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.  1    schematically illustrates an aircraft  10  that includes a pair of hybrid electric propulsion systems  100 A,  100 B (also referred to as hybrid gas turbine engines  100 A,  100 B or hybrid propulsion systems  100 A,  100 B). Each of the hybrid electric propulsion systems  100 A,  100 B includes a gas turbine engine  20  with a low speed spool  30  configured to drive rotation of a fan  42 . Gas turbine engine  20  also includes a high speed spool  32  that operates at higher speeds and pressures than the low speed spool  30 . A low spool motor  12 A is configured to augment rotational power of the low speed spool  30 . A high spool motor  12 B can be configured to augment rotational power of the high speed spool  32 . At least one power source  16  of the aircraft  10  can provide electrical power to the low spool motor  12 A and/or to the high spool motor  12 B. The power source  16  can be a stored energy source or a generator driven by an engine. For example, the power source  16  can include one or more of a battery, a super capacitor, an ultra capacitor, a fuel cell, a flywheel, and the like. Where the aircraft  10  includes an additional thermal engine (not depicted), such as an auxiliary power unit (APU), the power source  16  can be a generator driven by the thermal engine. Further, a generator of one of the hybrid electric propulsion systems  100 A,  100 B can provide power to the other hybrid electric propulsion systems  100 A,  100 B. For example, if the hybrid electric propulsion system  100 A is combusting fuel, the hybrid electric propulsion system  100 B may operate without burning fuel and can drive the low speed spool  30  based on the low spool motor  12 A receiving electric power from the hybrid electric propulsion system  100 A and/or the power source  16 . Further, the high speed spool  32  can be driven based on the high spool motor  12 B receiving electric power from the hybrid electric propulsion system  100 A and/or the power source  16 . 
     While the example of  FIG.  1    illustrates a simplified example of the gas turbine engine  20 , it will be understood that any number of spools, and inclusion or omission of 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. For example, rotor systems can be included in power generation systems, which may be ground-based as a fixed position or mobile system, and other such applications. 
       FIG.  2    illustrates a hybrid electric propulsion system  100  (also referred to as hybrid gas turbine engine  100  or hybrid propulsion system  100 ) as a further example of the hybrid electric propulsion system  100 A,  100 B of  FIG.  1   . In the example of  FIG.  2   , the hybrid electric propulsion system  100  includes gas turbine engine  20  operably coupled to an electrical power system  210  as part of a hybrid electric aircraft, such as aircraft  10  of  FIG.  1   . One or more mechanical power transmissions  150  (e.g.,  150 A,  150 B) can be operably coupled between the gas turbine engine  20  and the electrical power system  210 . The gas turbine engine  20  includes one or more spools, such as low speed spool  30  and high speed spool  32 , each with at least one compressor section and at least one turbine section operably coupled to a shaft (e.g., low pressure compressor  44  and low pressure turbine  46  coupled to inner shaft  40  and high pressure compressor  52  and high pressure turbine  54  coupled to outer shaft  50 ). The electrical power system  210  can include a low spool motor  12 A configured to augment rotational power of the low speed spool  30  and a high spool motor  12 B configured to augment rotational power of the high speed spool  32 . Although two motors  12 A,  12 B are depicted in  FIG.  2   , it will be understood that there may be only a single motor (e.g., only low spool motor  12 A) or additional motors (not depicted). Further, the motors  12 A,  12 B can be electric motors or alternate power sources may be used, such as hydraulic motors, pneumatic motors, and other such types of motors known in the art. The electrical power system  210  can also include a low spool generator  213 A configured to convert rotational power of the low speed spool  30  to electric power and a high spool generator  213 B configured to convert rotational power of the high speed spool  32  to electric power. Although two electric generators  213 A,  213 B (generally referred to as generators  213 A,  213 B) are depicted in  FIG.  2   , it will be understood that there may be only a single electric generator (e.g., only electric generator  213 B) or additional electric generators (not depicted). In some embodiments, one or more of the motors  12 A,  12 B can be configured as a motor or a generator depending upon an operational mode or system configuration, and thus one or more of the electric generators  213 A,  213 B may be omitted. 
     In the example of  FIG.  2   , the mechanical power transmission  150 A includes a gearbox operably coupled between the inner shaft  40  and a combination of the low spool motor  12 A and low spool generator  213 A. The mechanical power transmission  150 B can include a gearbox operably coupled between the outer shaft  50  and a combination of the high spool motor  12 B and high spool generator  213 B. In embodiments where the motors  12 A,  12 B are configurable between a motor and generator mode of operation, the mechanical power transmission  150 A,  150 B can include a clutch or other interfacing element(s). 
     The electrical power system  210  can also include motor drive electronics  214 A,  214 B operable to condition current to the motors  12 A,  12 B (e.g., DC-to-AC converters). The electrical power system  210  can also include rectifier electronics  215 A,  215 B operable to condition current from the electric generators  213 A,  213 B (e.g., AC-to-DC converters). The motor drive electronics  214 A,  214 B and rectifier electronics  215 A,  215 B can interface with an energy storage management system  216  that further interfaces with an energy storage system  218 . The energy storage management system  216  can be a bi-directional DC-DC converter that regulates voltages between energy storage system  218  and electronics  214 A,  214 B,  215 A,  215 B. The energy storage system  218  can include one or more energy storage devices, such as a battery, a super capacitor, an ultra capacitor, and the like. The energy storage management system  216  can facilitate various power transfers within the hybrid electric propulsion system  100 . The energy storage management system  216  may also transfer power to one or more electric motors on the engine, or to external loads  217  and receive power from one or more external power sources  219  (e.g., power source  16  of  FIG.  1   , aircraft power, auxiliary power unit power, cross-engine power, and the like). 
     A power conditioning unit  220  and/or other components can be powered by the energy storage system  218 . The power conditioning unit  220  can distribute electric power to support actuation and other functions of the gas turbine engine  20 . For example, the power conditioning unit  220  can power an integrated fuel control unit  222  to control fuel flow to the gas turbine engine  20 . The power conditioning unit  220  can also power a plurality of actuators (not depicted), such as bleed actuators, vane actuators, and the like. 
     One or more accessories  70  can also be driven by or otherwise interface with the gas turbine engine  20 . Examples of accessories  70  can include oil pumps, fuel pumps, and other such components. As one example, the accessories  70  include an oil pump driven through gearing, such as mechanical power transmission  150 B, in response to rotation of the high speed spool  32  and/or the high spool motor  12 B. Alternatively, accessories  70  can be electrically driven through power provided by the energy storage management system  216  or other such sources of electrical power. 
     Engagement and operation of the low spool motor  12 A, low spool generator  213 A, high spool motor  12 B, and high spool generator  213 B can change depending upon an operating state of the gas turbine engine  20  and any commands received. Collectively, any effectors that can change a state of the gas turbine engine  20  and/or the electrical power system  210  may be referred to as hybrid electric system control effectors  240 . Examples of the hybrid electric system control effectors  240  can include the motors  12 A,  12 B, electric generators  213 A,  213 B, integrated fuel control unit  222 , and/or other elements (not depicted). 
       FIG.  3    is a schematic diagram of control signal paths  250  of the hybrid electric propulsion system  100  of  FIG.  2    and is described with continued reference to  FIGS.  1  and  2   . A controller  256  can interface with the motor drive electronics  214 A,  214 B, rectifier electronics  215 A,  215 B, energy storage management system  216 , integrated fuel control unit  222 , accessories  70 , and/or other components (not depicted) of the hybrid electric propulsion system  100 . In embodiments, the controller  256  can control and monitor for fault conditions of the gas turbine engine  20  and/or the electrical power system  210 . For example, the controller  256  can be integrally formed or otherwise in communication with a full authority digital engine control (FADEC) of the gas turbine engine  20 . Alternatively, the controller  256  can be an aircraft level control or be distributed between one or more systems of the aircraft  10  of  FIG.  1   . In embodiments, the controller  256  can include a processing system  260 , a memory system  262 , and an input/output interface  264 . The controller  256  can also include various operational controls, such as a hybrid engine control  266  that controls the hybrid electric system control effectors  240  further described herein, for instance, based on a thrust command  270 . The thrust command  270  can be a throttle lever angle or a command derived based on a throttle lever angle control of the aircraft  10  of  FIG.  1   . 
     The processing system  260  can include any type or combination of central processing unit (CPU), including one or more of: a microprocessor, a digital signal processor (DSP), a microcontroller, an application specific integrated circuit (ASIC), a field programmable gate array (FPGA), or the like. The memory system  262  can store data and instructions that are executed by the processing system  260 . In embodiments, the memory system  262  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 algorithms in a non-transitory form. The input/output interface  264  is configured to collect sensor data from the one or more system sensors and interface with various components and subsystems, such as components of the motor drive electronics  214 A,  214 B, rectifier electronics  215 A,  215 B, energy storage management system  216 , integrated fuel control unit  222 , accessories  70 , and/or other components (not depicted) of the hybrid electric propulsion system  100 . The controller  256  provides a means for controlling the hybrid electric system control effectors  240  using a hybrid engine control  266  that can be dynamically updated during operation of the hybrid electric propulsion system  100 . The means for controlling the hybrid electric system control effectors  240  can be otherwise subdivided, distributed, or combined with other control elements. 
     The controller  256  with hybrid engine control  266  can apply control laws and access/update models to determine how to control and transfer power between the low speed spool  30  and high speed spool  32 . For example, sensed and/or derived parameters related to speed, flow rate, pressure ratios, temperature, thrust, and the like can be used to establish operational schedules and transition limits to maintain efficient operation of the gas turbine engine  20 . For instance, a mode of operation of the gas turbine engine  20 , such as idle, takeoff, climb, cruise, and descent can have different power settings, thrust requirements, flow requirements, and temperature effects. The hybrid engine control  266  can control electric current provided to the low spool motor  12 A and high spool motor  12 B and loading effects of the low spool generator  213 A and high spool generator  213 B. The hybrid engine control  266  can also determine a power split between delivering fuel to the combustor  56  and using the low spool motor  12 A and/or high spool motor  12 B to power rotation within the gas turbine engine  20 . 
     Referring now to  FIG.  4   , plot  300  graphically illustrates a relationship between engine spool speeds and time when transitioning through multiple operating modes. Line  302  indicates a percent speed  312  of the low speed spool  30  as time  310  advances and the hybrid electric propulsion system  100  transitions between e-taxi  306 , engine start  307 , and conventional idle  308 . E-taxi  306  refers to a mode of operation where the low spool motor  12 A drives rotation of the low speed spool  30  to produce thrust using the fan  42 , such that the aircraft  10  can be maneuvered on the ground without burning fuel in the combustor  56 . Line  304  indicates a percent speed  312  of the high speed spool  32  as time  310  advances and the hybrid electric propulsion system  100  transitions between e-taxi  306 , engine start  307 , and conventional idle  308 . As can be seen in  FIG.  4   , the high speed spool  32  can remain undriven during e-taxi mode  306 , which conserves energy by avoiding fuel burn and power draw from the high spool motor  12 B. In engine start  307 , the high spool motor  12 B can be used to increase the speed of the high speed spool  32  for light off and fuel burn in the combustor  56 . In conventional idle  308 , the motors  12 A,  12 B may not be needed, and the gas turbine engine  20  may be power by fuel burn. Alternatively, the engine-on idle state may include a further hybrid element where the idle state of the engine includes both fuel input and electric input to the electric motors  12 A,  12 B, or draw through the electric generators  213 A,  213 B. This is referred to as sub-idle, being possibly below conventional fuel-only idle in terms of either fuel flow and/or thrust. 
     Referring now to  FIG.  5   , plot  400  graphically illustrates a relationship between thrust  412  and throttle lever angle (TLA)  410 . Line  402  depicts an example thrust response starting at the e-taxi mode  306  of  FIG.  4   , where thrust  412  can be commanded below idle by controlling the low spool motor  12 A to drive rotation of the low speed spool  30  absent fuel burn in the combustor  56 . Generally, the operating mode of line  402  is for fuel off and electricity available as limited by a lower operating limit  403 . The lower operating limit  403  may be associated with a fuel-off detent of the TLA  410 . An idle level  407  may be associated with an idle detent of the TLA  410 . Line  404  depicts an example of a thrust response during engine start  307  of  FIG.  3   , where thrust  412  can be provided below an idle level  407  using the low spool motor  12 A to control thrust  412  while also using the high spool motor  12 B to control the high speed spool  32  to provide sufficient compression in the gas turbine engine  20  for light off in the combustor  56 . Line  406  depicts an example of a thrust response after starting the gas turbine engine  20  at idle level  407 , such as idle  308  of  FIG.  4   . Controlling the low spool motor  12 A and high spool motor  12 B can support a sub-idle operation state with thrust control at power settings lower than idle level  407 . Thrust  12  can be controlled at a demand and power output via the low spool motor  12 A and/or high spool motor  12 B for a thrust output less than a minimum thrust output at engine idle. The thrust response depicted at line  406  can start at idle level  407  and continue up in relation to TLA  410  along a response profile  408 . Although lines  402 ,  404 ,  406  and response profile  408  are depicted as substantially linear segments, it will be understood that lines  402 ,  404 ,  406  and response profile  408  can have other shapes and characteristics. 
       FIG.  5    further illustrates a first region  409  where the thrust response characteristic above the idle level  407  may be the same whether the fuel flow is on or off, and furthermore a second region  405  is defined below the idle level  407 . The similar thrust response characteristic can continue in the second region  405  to a lower thrust level before reaching the lower operating limit  403  at line  402 . A transition from the lower operating limit  403  to the idle level  407  can occur during engine start at line  404 . Line  404  is an example that can shift in position between lines  402  and  406  depending on the throttle lever angle  410  position for sub-idle operation. Power provided by the low spool motor  12 A and/or the high spool motor  12 B can support engine starting below idle level  407  within the second region  405 . 
     In embodiments, the controller  256  can blend the power distribution between the hybrid electric system control effectors  240  and fuel burn in the combustor  56 . From a pilot&#39;s perspective, the setting of throttle lever angle  410  produces thrust command  270  without the pilot having to distinguish between whether motor-based thrust or fuel burn based thrust is needed. While conventional systems may use detents to prevent a pilot from reducing thrust  412  below the idle level  407 , embodiments can support operation of thrust  412  down to line  402  to support e-taxi mode  306  and other intermediate modes of operation below conventional idle  308 . Thus, control of thrust  412  can be achieved before, during, and after engine start  307 . With respect to the aircraft  10 , the hybrid electric propulsion systems  100 A,  100 B can be independently controlled such that one of the hybrid electric propulsion systems  100 A,  100 B is operating in a fuel burning mode while the other of the hybrid electric propulsion systems  100 A,  100 B is operated using the low spool motor  12 A and/or the high spool motor  12 B or a blend of fuel burn and electric power. Such mixed modes of operation may be used, for instance, during descent of the aircraft  10 , where thrust  412  is desired from both gas turbine engines  20 , but only one of the gas turbine engines  20  actively burns fuel. Further, embodiments can support e-taxi mode  306  with warmup time to delay starting of the gas turbine engines  20  until reaching a location on the taxiway away from a boarding gate. 
     Referring now to  FIG.  6    with continued reference to  FIGS.  1 - 5   ,  FIG.  6    is a flow chart illustrating a method  600  for providing hybrid gas turbine engine starting control, in accordance with an embodiment. The method  600  may be performed, for example, by the hybrid electric propulsion system  100  of  FIG.  2   . For purposes of explanation, the method  600  is described primarily with respect to the hybrid electric propulsion system  100 ; however, it will be understood that the method  600  can be performed on other configurations (not depicted). 
     Method  600  pertains to the controller  256  executing embedded code for the starting and thrust control using hybrid engine control  266  along with other control functions. At block  602 , the controller  256  can receive a thrust command  270  for a gas turbine engine  20 , where the gas turbine engine  20  includes a low speed spool  30 , a high speed spool  32 , and a combustor  56 . The controller  256  is configured to cause fuel flow to the combustor  56  under certain operating conditions. 
     At block  604 , the controller  256  can control a low spool motor  12 A to drive rotation of the low speed spool  30  responsive to the thrust command  270  while the controller  256  does not command fuel flow to the combustor  56 , where the low spool motor  12 A is configured to augment rotational power of the low speed spool  30 . Fuel flow can be reduced or completely shut off depending upon an operating state of the gas turbine engine  20 . For example, the controller  256  can output a command of no fuel, fuel flow off, and/or otherwise effectively disable or reduce fuel flow as targeted. The operating state can depend on a combination of commands, conditions, and modes, such as an e-taxi mode, a starting mode, a ground idle mode, a takeoff mode, a climb mode, a cruise mode, an in-flight idle mode, a descent mode, a landing mode, and other such modes. The controller  256  can determine an allocation of the thrust command  270  between commanding fuel flow to the combustor  56  and electric current to the low spool motor  12 A based on the operating state of the gas turbine engine  20  and a throttle lever angle  410 , where the throttle lever angle  410  can be received from a pilot control, an auto-pilot control, or other such source on the aircraft  10 . The low spool motor  12 A can be powered by one or more of a generator, an energy storage system, and a power source  16  external to the gas turbine engine  20 . 
     At block  606 , the controller  256  can control the low spool motor  12 A responsive to the thrust command  270  during a starting operation of the gas turbine engine  20 . The starting operation can be a ground-based start or an in-flight restart. 
     At block  608 , the controller  256  can control the low spool motor  12 A to drive rotation of the low speed spool  30  responsive to the thrust command at or above an idle condition of the gas turbine engine  20 . 
     In some embodiments, a high spool motor  12 B can be used in conjunction with the low spool motor  12 A. For example, the controller  256  can receive an engine start command  610 . At block  612 , the controller  256  can control a high spool motor  12 B to accelerate the high speed spool  32  responsive to a start command while the low spool motor  12 A controls thrust of the gas turbine engine  20  on the low speed spool  30 , where the high spool motor  12 B is configured to augment rotational power of the high speed spool  32 . Control of the high spool motor  12 B of block  612  can occur in parallel with control of the low spool motor  12 A of block  604  or blocks  604  and  612  can be other sequenced, combined, or further subdivided. The controller  256  can be configured to control a thrust response of the gas turbine engine  20  to a response profile  408  based on the throttle lever angle  410  using any combination of the low spool motor  12 A, high spool motor  12 B, and fuel burn. 
     In some embodiments, a low spool generator  213 A is configured to extract power from the low speed spool  30 , and a high spool generator  213 B is configured to extract power from the high speed spool  32 . The controller  256  can be configured to selectively provide electrical power from the low spool generator  213 A to the high spool motor  12 B and selectively provide electrical power from the high spool generator  213 B to the low spool motor  12 A. The controller  256  can also be configured to selectively engage either or both of the low spool generator  213 A and the high spool generator  213 B to adjust a load and speed of either or both of the low speed spool  30  and the high speed spool  32 . 
     While the above description has described the flow process of  FIG.  6    in 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. Also, it is clear to one of ordinary skill in the art that, the starting control described herein can be combined with and enhance other control features, such as valves, vanes, and fuel flow control. 
     Although some embodiments described herein relate to relighting an engine during or after an e-taxi event, it should be appreciated that the disclosed techniques for relighting an engine can also apply to other modes of operation and/or flight phases. For example, a mode of operation of the gas turbine engine  20  of  FIGS.  1  and  2   , such as idle, takeoff, climb, cruise, and descent can have different power settings, thrust requirements, flow requirements, and temperature effects. 
     An aircraft can selectively power a hybrid electric engine by providing electric power from various sources, such as a battery system, another engine, and/or an APU or secondary power unit (SPU). With respect to the aircraft  10  of  FIG.  1   , the hybrid electric propulsion systems  100 A,  100 B can be independently controlled such that one of the hybrid electric propulsion systems  100 A,  100 B is operating in a fuel burning mode while the other of the hybrid electric propulsion systems  100 A,  100 B is operated using the low spool motor  12 A and/or the high spool motor  12 B or a blend of fuel burn and electric power. Such mixed modes of operation may be used, for instance, during descent of the aircraft  10 , where thrust is desired from both gas turbine engines  20 , but only one of the gas turbine engines  20  actively burns fuel. In such cases, it may be desirable, during descent, to restart (e.g., relight) the one of the gas turbine engines  20  that is not actively burning fuel. 
     During descent with one of the engines  20  operating on electric power and the other of the engines  20  operating with fuel burn, various aspects may be considered such that the engine system operates efficiently, and the one of the engines  20  operating on electric power can rapidly resume a fuel-burn mode of operation (e.g., restart or relight). However, in some cases, it may be desirable, during descent, to keep only one of the gas turbine engines  20  actively burning fuel while the other of the gas turbine engines  20  is operating on electric power. The choice of whether to use only one of the gas turbine engines  20  in fuel burning mode or to use both of the gas turbine engines  20  in fuel burning mode can depend, for example, on the descent angle (e.g., glide slope). 
     One or more embodiments described herein provides a process for determining when to enable a hybrid electric engine to operate in a fuel burning mode or to operate in an electrically powered mode during descent. For example, in some situations such as e-taxi or descent, an aircraft having two hybrid electric engines can operate with one of the engines in fuel-burning mode and the other engine operating in an electric power (non-fuel burning) mode. This can be referred to as single-engine descent when the aircraft is in a descent phase of flight. 
     During single engine descent, thrust matching is performed. For example, 1000 pounds of thrust per each of the engines  20  can be achieved by driving a fan of one of the engines  20  electrically to provide 1000 pounds of thrust, and the other of the engines  20 , operating in a fuel-burning mode, can generate power for the electrically-operated engine and produce 1000 pounds of thrust. 
     In some cases, in some cases of descent, such as with a relatively shallower glide slope, it is more fuel efficient to operate both engines  20  in the fuel-burning mode. However, in cases with a relatively steeper glide slope, it is more fuel efficient to maintain a single engine descent (e.g., to operate only one of the engines  20  in the fuel-burning mode while the other engine  20  operates in the electrically powered mode). 
     A control strategy implemented by the hybrid electric engine system can consider a planned descent profile (e.g., a desired glide slope), a battery state of charge, and/or an estimate of transmission losses to decide whether to exercise this function of single engine descent. Other parameters may also be considered and can include, for example, the availability of power from an APU/SPU as a primary or backup power source for the electrically operated engine. 
     With a steep approach or continuous approach, it may make more sense to operate in single engine descent mode, which supports a steeper approach than a traditional duel fuel-driven engine approach. Glide range to alternate landing locations may also be considered as an input to activating single engine descent mode. For example, if the glide range to an alternate landing location is determined to exceed a distance threshold, the mode of the electrically powered engine can be changed to the fuel-burning mode. As an example, the distance threshold could be present, could be based on factors such as fuel burn rate, available fuel, environmental factors, pilot input, and the like, including combinations thereof. 
     According to one or more embodiments described herein, a tradeoff of expected fuel savings versus component wear effects can be determined prior to activating single engine descent mode. For example, where fuel savings is greater than any negative effects on component wear, it may be desirable to use single engine descent. However, where negative effects on component wear is greater, it may be desirable to use a traditional duel fuel-driven engine approach. Flight time with single engine descent mode active can be tracked for each engine as well as whether each engine operated as a fuel-driven engine or an electrically operated engine during single engine descent mode. This information can be used to determine lifetime estimates for engines/components. 
       FIG.  7    is a flow chart illustrating a method  700 , in accordance with an embodiment of the disclosure. The method  700  can be performed by any suitable system, device, controller, etc., such as the controller  256 . It should be appreciated that system, device, controller, etc., that performs the method can be located in a control unit of one of the engines  20 , both of the engines  20 , the aircraft  10 , combinations thereof, or at another location. 
     At block  702 , a controller (e.g., the controller  256 ) determines a thrust requirement to satisfy a desired glide slope. The desired slope can be received, for example, from a system/controller of the aircraft, from the pilot, from a remote ground-based system, and/or the like. The desired glide slope indicates a desired path of descent of the aircraft preparing to land. Some landing locations (e.g., airports) require relatively higher glide slopes than other landing locations due to geographic conditions, environmental conditions, local regulations, etc. Further, some aircraft operate more efficiently at certain glide slopes than others. 
     At decision block  704 , the controller determines, based on the thrust requirement, whether thrust matching can be maintained while operating a first gas turbine engine (e.g., one of the engines  20 ) in a fuel-burning mode and operating a second gas turbine engine (e.g., the other of the engines  20 ) in an electrically powered mode. As described herein, during single engine descent, thrust matching is performed. For example, 1000 pounds of thrust per each of the engines  20  can be achieved by driving a fan of one of the engines  20  electrically to provide 1000 pounds of thrust, and the other of the engines  20 , operating in a fuel-burning mode, can generate power for the electrically-operated engine and produce 1000 pounds of thrust. The controller determines whether the thrust matching can be maintained while the aircraft operates in single engine descent. Determining whether thrust matching can be maintained can include comparing a first thrust of the first engine with a second thrust of the second engine. It is determined that thrust matching can be maintained when the second thrust satisfies a threshold difference relative to the first thrust. However, it is determined that thrust matching cannot be maintained when the second thrust fails to satisfy a threshold difference relative to the first thrust. The threshold difference could be a percent difference between the first and second thrusts (e.g., 2% difference), an absolute value difference between the first and second thrusts (e.g., 10 pounds of thrust difference), and/or the like. 
     If, at decision block  704 , it is determined that thrust matching cannot be maintained, the controller, at block  706 , commands fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. That is, the second engine ceases to operate in the electrically powered mode and switches to fuel-burning mode, thus providing the aircraft with dual engine descent. This ensures that the thrust matching can be maintained while maintaining the desired glide slope. 
     If, at decision block  704 , it is determined that thrust matching can be maintained, at block  708 , the first gas turbine engine continues to operate in the fuel-burning mode and the second gas turbine engine continues to operate in the electrically powered mode. 
     Additional processes also may be included. For example, the method  700  can include determining a glide range to an alternate landing location and then, responsive to determining that the glide range exceeds a distance threshold, commanding fuel flow to a combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode. 
     According to one or more embodiments described herein, the method  700  includes tracking a first amount of time the second engine spends in the fuel-burning mode and a second amount of time the second engine spends in the electrically powered mode. The commanding can then be based on the first and/or second amount of time. For example, commanding the fuel flow to the combustor of the second engine to cause the second gas turbine engine to operate in the fuel-burning mode is based at least in part on at least one of the first amount of time or the second amount of time. 
     It should be understood that the process depicted in  FIG.  7    represents an illustration, and that other processes may be added or existing processes may be removed, modified, or rearranged without departing from the scope of the present disclosure. 
     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.