Patent Publication Number: US-2023139529-A1

Title: Hybrid-electric single engine descent failure management

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 63/273,452 filed Oct. 29, 2021, the contents of which are hereby incorporated by reference in its entirety. 
    
    
     BACKGROUND 
     Exemplary embodiments pertain to the art of gas turbine engines, and in particular to aircraft hybrid-electric engines. 
     An aircraft can selectively power a hybrid-electric engine by providing electric power from various sources. During descent with one engine turned off with thrust provided by electric power and the other engine operating with fuel burn, various aspects must be considered such that the engine system operates efficiently, and such that the engine operating on electric power can rapidly resume a fuel-burn mode of operation if and when that is needed. 
     For example, during single engine descent, thrust matching is performed. For example, 1000 pounds of thrust per engine would be achieved by driving one engine fan electrically to provide 1000 pounds of thrust and by driving the other engine in a fuel-burning mode can generate power for the electrically-operated engine and produce 1000 pounds of thrust. 
     BRIEF DESCRIPTION 
     According to an aspect of the disclosure, a hybrid-electric aircraft system is provided and includes first and second hybrid-electric engines, each of which includes an electric motor to drive operations thereof, and a supplemental power unit (SPU). The SPU is configured as a thermal engine paired with a generator and is configured to generate electrical power. The first and second hybrid-electric engines are operable normally and off, respectively, with electrical power generated by the SPU being diverted to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the second hybrid-electric engine is off during descent. 
     In accordance with additional or alternative embodiments, normal operation of the first hybrid-electric engine includes the first hybrid-electric engine operating at a minimum descent idle level. 
     In accordance with additional or alternative embodiments, the second hybrid-electric engine being off is simultaneous with diversion of the electrical power generated by the SPU to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the SPU includes a turbine operably disposed within an aircraft fuselage. 
     In accordance with additional or alternative embodiments, the SPU includes an SPU controller configured to independently operate the SPU. 
     In accordance with additional or alternative embodiments, each of the first and second hybrid-electric engines includes a full authority digital engine control (FADEC) and the SPU includes an SPU controller which is configured to independently operate the SPU and which is directly communicative with the FADEC of each of the first and second hybrid-electric engines. 
     In accordance with additional or alternative embodiments, the electrical power generated by the SPU is blended with electrical power from one or more other sources for diversion to the electric motor of the second hybrid-electric engine. 
     According to an aspect of the disclosure, a hybrid-electric aircraft system is provided and includes first and second hybrid-electric engines, each of which includes an electric motor to drive operations thereof, a supplemental power unit (SPU) configured as a thermal engine paired with a generator and to generate electrical power and a control system. The control system is operably coupled to each of the first and second hybrid-electric engines and to the SPU. The control system is configured to run the first hybrid-electric engine normally with the second hybrid-electric engine off and to divert electrical power generated by the SPU to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the control system diverts electrical power generated by the SPU to the electric motor of the second hybrid-electric engine during descent. 
     In accordance with additional or alternative embodiments, normal running of the first hybrid-electric engine includes running the first hybrid-electric engine at a minimum descent idle level. 
     In accordance with additional or alternative embodiments, the second hybrid-electric engine being off is simultaneous with electrical power generated by the SPU being diverted to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the SPU includes a turbine operably disposed within an aircraft fuselage. 
     In accordance with additional or alternative embodiments, the SPU includes an SPU controller configured to operate the SPU independently from the control system. 
     In accordance with additional or alternative embodiments, each of the first and second hybrid-electric engines includes a full authority digital engine control (FADEC) and the SPU includes an SPU controller which is configured to operate the SPU independently from the control system and which is directly communicative with the FADEC of each of the first and second hybrid-electric engines. 
     In accordance with additional or alternative embodiments, each of the first and second hybrid-electric engines includes a full authority digital engine control (FADEC), the SPU includes an SPU controller and the control system is communicative with the FADEC of each of the first and second hybrid-electric engines and the SPU controller. 
     In accordance with additional or alternative embodiments, the control system is further configured to blend electrical power generated by the SPU with electrical power from one or more other sources for diversion to the electric motor of the second hybrid-electric engine. 
     According to as aspect of the disclosure, a method of operating a hybrid-electric aircraft system including first and second hybrid-electric engines, each of which includes an electric motor, and a supplemental power unit (SPU) configured to generate electrical power is provided. The method includes initiating a fuel conservation mode, running the first hybrid-electric engine normally with the second hybrid-electric engine being off and diverting electrical power generated by the SPU to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the second hybrid-electric engine being off is simultaneous with the diverting of the electrical power generated by the SPU to the electric motor of the second hybrid-electric engine. 
     In accordance with additional or alternative embodiments, the second hybrid-electric engine being off and the diverting of the electrical power generated by the SPU to the electric motor of the second hybrid-electric engine are executed during descent. 
     In accordance with additional or alternative embodiments, the running of the first and second hybrid-electric engines normally and in the low power mode, respectively, and the directing of bleed air from the first hybrid-electric engine to the second hybrid-electric engine are simultaneous. 
    
    
     
       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    a partial cross-sectional illustration of a gas turbine engine according to a non-limiting embodiment; 
         FIG.  2    is a schematic diagram of a hybrid-electric gas turbine engine system including a power management system in accordance with a non-limiting embodiment of the disclosure; 
         FIG.  3    is a block diagram of a hybrid-electric engine management system according to a non-limiting embodiment; 
         FIG.  4    is a block diagram illustrating a hybrid-electric aircraft system in accordance with embodiments; 
         FIG.  5    is a schematic diagram illustrating communicative components of the hybrid-electric aircraft system of  FIG.  4    in accordance wither embodiments; and 
         FIG.  6    is a flow diagram illustrating a method of operating a hybrid-electric aircraft system in accordance with embodiments. 
     
    
    
     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. 
     Electric generators and motors have been incorporated into aircraft engines to provide a hybrid-electric engine that produces electric energy for various engine and aircraft support systems. However, primary power production and operation relies on conversion of the high-energy exhaust gas flow into mechanical power. Aircraft control systems for hybrid-electric engines systems have been developed, which allow for selectively powering the hybrid-electric engine by providing electric power from various sources to reduce fuel consumption and improve overall engine efficiencies. During descent with one engine operating on electric power and the other engine operating with fuel burn, various aspects must be considered such that the engine system operates efficiently, and the engine operating on electric power can rapidly resume a fuel-burn mode of operation. 
     With reference now to  FIG.  1   , a gas turbine engine  20  is illustrated according to a non-limiting embodiment. The gas turbine engine  20  is disclosed herein as a multi-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . Alternative engines might include other systems or features. The fan section  22  drives air along a bypass flow path B in a bypass duct, while the compressor section  24  drives air along a core flow path C for compression and communication into the combustor section  26  then expansion through the turbine section  28 . 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 multi-spool turbofans as the teachings may be applied to other types of turbine engines including, for example, three-spool architectures. 
     The exemplary engine  20  generally includes one or more low-spool generator machines  30 , referred to herein as a “low-spool”  30  and a high-spool generator machine  32 , referred to herein as a “high-spool  32 ” mounted for rotation about an engine central longitudinal axis (A) relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low-spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low-pressure compressor  44  and a low-pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low-spool  30 . The high-spool  32  includes an outer shaft  50  that interconnects a high-pressure compressor  52  and high-pressure turbine  54 . A combustor  56  is arranged in exemplary gas turbine  20  between the high-pressure compressor  52  and the high-pressure turbine  54 . An engine static structure  36  is arranged generally between the high-pressure turbine  54  and the low-pressure turbine  46 . The engine static structure  36  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     The core airflow is compressed by the low-pressure compressor  44  then the high-pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded over the high-pressure turbine  54  and low-pressure turbine  46 . The turbines  46 ,  54  rotationally drive a respective low-spool  30  and high-spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of combustor section  26  or even aft of turbine section  28 , and fan section  22  may be positioned forward or aft of the location of gear system  48 . 
     The engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6), with an example embodiment being greater than about ten (10), the geared architecture  48  is 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 turbine  46  has a pressure ratio that is greater than about five. In one disclosed embodiment, the engine  20  bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low-pressure compressor  44 , and the low-pressure turbine  46  has a pressure ratio that is greater than about five 5:1. Low-pressure turbine  46  pressure ratio is pressure measured prior to inlet of low-pressure turbine  46  as related to the pressure at the outlet of the low-pressure turbine  46  prior to an exhaust nozzle. The geared architecture  48  may 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 section  22  of the engine  20  is 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. In one or more non-limiting embodiments, the temperature correction can be determined according to the relationship of [(Tram ° R)/(518.7° R)]0.5, where Tram is a ram air temperature. 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 of  FIG.  1    illustrates one example of the gas turbine engine  20 , it will be understood that any number of spools, inclusion or omission of the gear system  48 , 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. 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 portions of a hybrid-electric gas turbine engine system  200  according to a non-limiting embodiment. The hybrid-electric gas turbine engine system  200  includes a rotor system  202  and a power management system  205 . The rotor system  202  includes at least one compressor section  204  and at least one turbine section  208  operably coupled to a shaft  206 . The rotor system  202  can include one or more spools of the gas turbine engine  20  of  FIG.  1   , such as one or more low-spools  30  and/or one or more high-spools  32 . For example, employing one or more low-spools  30 , one or more compressor sections  204  can be equivalent to the low-pressure compressor  44 , the shaft  206  can be equivalent to the inner shaft  40 , and the at least one turbine section  208  can be equivalent to the low-pressure turbine  46  of  FIG.  1   . When embodied as the high-spool  32 , one or more compressor sections  204  can be equivalent to the high-pressure compressor  52 , the shaft  206  can be equivalent to the outer shaft  50 , and the at least one turbine section  208  can be equivalent to the high-pressure turbine  54  of  FIG.  1   . 
     In the example of  FIG.  2   , a battery charging system  210  is operably coupled to the rotor system  202 . The battery charging system  210  includes a generator  212  operably coupled to the shaft  206 . In the example of  FIG.  2   , a geared interface  230  operably couples the generator  212  to the shaft  206 . The geared interface  230  can include, for instance, an auxiliary gear  233  coupled to an auxiliary shaft  235  driven by the generator  212 . The geared interface  230  can also include a rotor gear  237  coupled to the shaft  206 . The auxiliary gear  233  and the rotor gear  237  can each be beveled gears. The auxiliary shaft  235  can be a tower shaft that enables the generator  212  to be separated at a greater distance from the rotor system  202  than direct coupling to the shaft  206  would provide. Further separation of the generator  212  from the rotor system  202  can improve accessibility to the generator  212  for servicing and may reduce heating effects of the rotor system  202  on the generator  212  (e.g., due to fuel combustion). A disconnect  240 , such as a clutch, can be positioned between the generator  212  and a portion of the shaft  206  such that the generator  212  can be selectively engaged and disengaged to rotate with rotation of the shaft  206 . In alternate embodiments, the generator  212  is operably coupled to the shaft  206  absent the geared interface  230  (e.g., direct coupling). 
     The battery charging system  210  also includes a power converter system  214  in signal communication with the generator  212  and a battery system  255 . In some embodiments, the generator  212  is a motor-generator configurable in a generator mode to charge a rechargeable battery included in the battery system, and in a motor mode to provide supplemental rotation force to the rotor system  202  of gas turbine engine  20  of  FIG.  1   . The power converter system  214  includes converter electronics configured to condition current from the generator  212  such that the battery included in the battery system  255  can be repeatedly recharged. The converter electronics include, but are not limited to, analog current (AC) distribution circuitry, bi-directional power electronic circuitry, direct current (DC) power distribution electronics, AC-to-DC converter electronics, DC-to-DC converter electronics, rectifier circuits, a battery system, and an auxiliary power unit (APU)/supplemental power unit (SPU). 
     The generator  212  can include conventional generator/motor components, such as a rotor and stator, including a plurality of windings and/or permanent magnets. The converter electronics  214  can also include conventional current control electronics, such as filters, switching components, rectifiers, inverters, voltage converters, and the like. The generator  212  can perform as a variable frequency generator in a generator mode due to speed fluctuations of rotation of the shaft  206 , which may be primarily driven by the at least one turbine section  208 . Alternatively, a frequency normalizing component can interface with the generator  212  to produce a constant frequency output (e.g., through the converter electronics  214  or as a mechanical interface between the generator  212  and the shaft  206 ). In some embodiments, the generator  212  may be operable as a starter motor to partially or completely power rotation of the shaft  206  in a starting mode of operation (e.g., to start the gas turbine engine  20  of  FIG.  1   ) and/or can provide supplemental power to the shaft  206  during various flight phases of the hybrid-electric aircraft  200 . Other uses and functions for the generator  212  are contemplated. 
     The converter electronics  214  can control charging of the battery system  255  responsive to a controller  216 . The controller  216  can enable a flow of a charging current from the generator  212  or a power input  252  to charge the battery included in the battery system  255  as regulated and conditioned through the converter electronics  214 . The power input  252  can be an external input, such as power received through a plug interface while the hybrid-electric aircraft  200  is on the ground at a ground-based power source, e.g., at a gate or service location. In some embodiments, the converter electronics  214  may receive electric current from an auxiliary power input  254  to provide a supplemental or alternative power source for charging the battery included in the battery system  255 . For instance, the auxiliary power input  254  may receive electric current from an auxiliary power unit (not depicted) or another instance of the gas turbine engine  20  on the hybrid-electric aircraft  200 . The charge stored in the battery system  255  can provide an electric current for a propulsion system use  256 , which may include powering one or more electric motors of the hybrid-electric aircraft  200  during various operational states and/or providing power to the generator  212  when operating in a motor mode, for instance, to assist in driving rotation of shaft  206 . The propulsion system uses  256  can be part of the gas turbine engine  20  that includes the rotor system  202  or another aircraft system, such as another instance of the gas turbine engine  20  on the hybrid-electric aircraft  200 . 
     In embodiments, the controller  216  of the battery charging system  210  can monitor one or more rotor system sensors  218  while the rotor system  202  is rotating. The rotor system sensors  218  can be any type or combination of sensors operable to measure aspects of the motion of the rotor system  202 . For example, the rotor system sensors  218  can include one or more accelerometers, speed sensors, torque sensors, and the like. The rotor system sensors  218  can include existing sensors used for controlling the gas turbine engine  20 . The controller  216  can control a charging of the battery system  255 , for instance, by selecting the source of electric current received through the converter electronics  214 . Data collected from the rotor system sensors  218  can be used to determine an operational status of a gas turbine engine  20  of  FIG.  2   . Alternatively, the operational status of a gas turbine engine  20  can be received as a signal or message from an alternate source, such as an engine system or aircraft communication bus. The controller  216  may also control other system aspects, such as controlling operation of the gas turbine engine  20  of  FIG.  1   . For example, the controller  216  can be integrally formed or otherwise in communication with a full authority digital engine control (FADEC) of the gas turbine engine  20 . The rotor system sensors  218  need not be directly coupled to the controller  216 , as sensor data or sensor-derived data can be observed or determined by another control (e.g., a FADEC) and provided to the controller  216 . 
     In embodiments, the controller  216  can include a processing system  220 , a memory system  222 , and an input/output interface  224 . The processing system  220  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  222  can store data and instructions that are executed by the processing system  220 . In embodiments, the memory system  222  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  224  is configured to collect sensor data from the one or more rotor system sensors  218  and interface with the power converter system  214  and/or other systems (not depicted). 
     Turning now to  FIG.  3   , a hybrid-electric engine management system  300  is illustrated. The hybrid-electric engine management system  300  includes a first hybrid-electric engine system  200   a , a second hybrid-electric engine syst  200   b , a power converter system  214 , an aircraft power system  310 , and an aircraft controller  216 . 
     The first hybrid-electric engine system  200   a  includes a first gas turbine engine  20   a , a low-spool electric machine  30   a , an optional second low-spool electric machine (not shown), and a high-spool electric machine  32   a . The high-spool electric machine  32   a  can be coupled to the first hybrid-electric engine system  20   a  via a geared interface  232   a . Similarly, the electric machine  30   a  is coupled to the first gas turbine engine  20   a  via a geared interface  230   a.    
     The second hybrid-electric engine system  200   b  includes a second gas turbine engine  20   b , a low-spool electric machine  30   b , an optional second low-spool electric machine (not shown), and a high-spool electric machine  32   b . The high-spool electric machine  32   b  is coupled to the second hybrid-electric engine system  20   b  via a geared interface  232   b . Similarly, the low-spool electric machine  30   b  can be coupled to the second gas turbine engine  20   b  via a geared interface  230   b.    
     The power converter system  214  includes a direct current (DC) power distribution bus  320 . The DC power distribution bus  320  includes multiple contactors and switches by which the DC power distribution bus  320  can connect to the low-spool electric machine  30   a  via bi-directional power electronics  321  and an alternating current (AC) distribution bus  322 , to the low-spool electric machine  30   b  via bi-directional power electronics  323  and an AC distribution bus  324 , to the high-spool electric machine  32   a  via bi-directional power electronics  325  and an alternating current (AC) bus  326  and to the high-spool electric machine  32   b  via bi-directional power electronics  327  and an AC distribution bus  328 . In addition, the DC power distribution bus  320  also includes multiple contactors and switches by which the DC power distribution bus  320  can connect to the battery system  255  via a DC/DC converter  329  and to supplemental power unit (SPU)  330  via rectifier  331 . 
     The SPU  330  can be configured as a turbine engine that is normally disposed in an aft section of a tail of an aircraft and will be described in further detail below. 
     Thus, the power converter system  214  includes converter electronics configured to provide power to the aircraft power system  310  and/or the battery system  255 . For example, the power converter system  214  can deliver power to the aircraft power system  310 , which can be utilized to power the aircraft galley, cabin lighting system, cabin HVAC systems, etc. The power converter system  214  can also deliver power to the battery system  255  to charge one or more rechargeable batteries. The power converter system  214  can also deliver power from the battery system  255  to the aircraft power system  310  and/or to the first and or second hybrid engine system  200   a  and  200   b.    
     The controller  216  is in signal communication with the first hybrid-electric engine system  200   a , the second hybrid-electric engine system  200   b , the power converter system  214  and the aircraft power system  310 . In one or more non-limiting embodiments, the controller  216  can selectively control various operations of the first and second hybrid-electric engine systems  200   a  and  200   b  such as, for example, selectively activating and deactivating the first and/or second engines  20   a  and  20   b . The controller  216  can also selectively activate and deactivate the low-spool electric machine  30   a  and the high-spool electric machine  32   a  included in the first hybrid-electric engine system  200   a  and/or the low-spool electric machine  30   b  and the high-spool electric machine  32   b  included in the second hybrid-electric engine systems  200   b . In one or more non-limiting embodiments, the controller can selectively activate the low-spool electric machines and the high-spool electric machine independently from one another. In one or more non-limiting embodiments, the controller  216  can selectively activate or deactivate one of the low-spool electric machines (e.g., low-spool electric machine  31   a ) included in a given hybrid-electric engine systems with respect to other low-spool electric machines (e.g., low-spool electric machine  30   a ) included in a given hybrid-electric engine system (e.g., hybrid-electric engine system  200   a ). 
     With continued reference to  FIG.  3   , to reduce fuel consumption and improve overall operating efficiency during descent of the aircraft, the aircraft controller  216  can deactivate the second gas turbine engine  20   b  while maintaining activation of the first gas turbine engine  20   a . Accordingly, power generated by the low-spool electric machine  30   a  included in the first hybrid-electric engine system  200   a  is conditioned by the power converter system  214 , and delivered to the low-spool electric machine  30   b  included in the second hybrid-electric engine system  200   b . Alternatively, to reduce fuel consumption and improve overall operating efficiency during descent of the aircraft but without deactivating the second gas turbine engine  20   b , the aircraft controller  216  can control an operation of the second gas turbine engine  20   b  to operate at low power levels and with minimal fuel burn while maintaining activation of the first gas turbine engine  20   a . Maintaining the second gas turbine engine  20   b  in a lit condition will increase reliability, for instance, if there is an electrical fault or if descent is aborted. 
     Thus, with continued reference to  FIG.  3    and with additional reference to  FIG.  4   , a hybrid-electric aircraft system  401  is provided to allow the second gas turbine engine  20   b  (see  FIG.  3   ) to be operated at the low power levels and with minimal fuel burn or turned off within no fuel burn during certain flight envelopes such as, but not limited to, descent for example. The hybrid-electric aircraft system  401  can be deployed in an aircraft  402  with first and second wings  403  and  404  that extend outwardly from fuselage  405  having a tail section  406 . The hybrid-electric aircraft system  401  includes a first hybrid-electric engine  410  arranged on the first wing  403 , a second hybrid-electric engine  420  arranged on the second wing  404 , an SPU  430  and a control system (i.e., the controller  216  of  FIG.  3   ). The hybrid-electric aircraft system  401  can further include additional or alternative power sources  440  (e.g., the battery system  255  of  FIG.  3    or an auxiliary power unit or APU). 
     The first hybrid-electric engine  410  includes a first electric motor  411  that is configured to drive certain operations of the first hybrid-electric engine  410  and the second hybrid-electric engine  420  includes a second electric motor  421  that is configured to drive certain operations of the second hybrid-electric engine  420  (see the description of the first and second hybrid-electric engine systems  200   a  and  200   b  provided above with reference to  FIG.  3   ). The SPU  430  can include or be provided as a turbine that is operably disposed within the fuselage  405  or, in some cases, the tail section  406  of the fuselage  405 . The SPU  430  is configured as a thermal engine paired with a generator and to generate electrical power that can be distributed to various systems on the aircraft  402  including, but not limited, to either the first electric motor  411  or the second electric motor  421 . 
     In accordance with embodiments, in order to conserve fuel for example during descent, the first hybrid-electric engine  410  can run normally or at a minimum descent idle level, the second hybrid-electric engine  420  can run in a low-power mode or below the minimum descent idle level or be turned off and the electrical power generated by the SPU  430  can be diverted to the second electric motor  421  of the second hybrid-electric engine  420  to power the second electric motor  421  to continue to drive certain non-thermal, rotational operations of the second hybrid-electric engine  420 . In these or other cases, the electrical power generated by the SPU  430  can be blended with electrical power drawn from the additional or alternative power sources  440 . The running of the first hybrid-electric engine  410  normally or at the minimum descent idle level, the running of the second hybrid-electric engine  420  in the low-power mode or below the minimum descent idle level or the second hybrid-electric engine  420  being turned off can be executed simultaneously with the diversion of the electrical power generated by the SPU  430  to the second electric motor  421 . 
     In accordance with further embodiments, in order to conserve fuel for example during descent, the control system or the controller  216  of  FIG.  3    can be operably coupled to each of the first and second hybrid-electric engines  410  and  420  and to the SPU  430 . In this condition, the control system can be configured to run or operate the first hybrid-electric engine  410  normally or at the minimum descent idle level, to run or operate the second hybrid-electric engine  420  in the low-power mode or below the minimum descent idle level or to turn the second hybrid-electric engine  42  off and to divert the electrical power generated by the SPU  430  to the second electric motor  421  of the second hybrid-electric engine  420  to power the second electric motor  421  to continue to drive certain non-thermal, rotational operations of the second hybrid-electric engine  420 . In these or other cases, the control system can blend the electrical power generated by the SPU  430  with the electrical power drawn from the additional or alternative power sources  440 . The running of the first hybrid-electric engine  410  normally or at the minimum descent idle level and the running of the second hybrid-electric engine  42  in the low-power mode or below the minimum descent idle level or the turning of the second hybrid-electric engine  42  off can be executed by the control system simultaneously with the diversion of the electrical power generated by the SPU  430  to the second electric motor  421 . 
     With reference to  FIG.  5   , the SPU  430  can include an SPU controller  431  that is configured to operate the SPU  430  independently from the control system or the controller  216  of  FIG.  3   . That is, the SPU controller  431  can cause the SPU  430  to spool up automatically or based on the SPU controller  431  determining that the SPU  430  should be operational. In accordance with additional or alternative embodiments, each of the first and second hybrid-electric engines  410  and  420  can include a full authority digital engine control (FADEC)  412  and  422  and, in these or other cases, the SPU controller  431  can be directly communicative with the FADEC  412  and the FADEC  422 . By way of such communications, the SPU controller  431  can independently determine an operational condition of either or both of the first and second hybrid-electric engines  410  and  420  and thus cause the SPU  430  to spool up accordingly. In addition, the SPU controller  431  can be configured to determine whether communications with either of the FADECs  412  and  422  is inactive or otherwise indicative of a fault and thus cause the SPU  430  to spool up accordingly. In accordance with still further additional or alternative embodiments, the FADECs  412  and  422  and the SPU controller  431  can each be communicative with the control system or the controller  216  of  FIG.  3   . In these or other cases, the SPU controller  431  can control operations of the SPU  430  in accordance with commands provided by the control system or the controller  216  of  FIG.  3   . 
     As shown in  FIG.  5   , the SPU controller  431  can include or be provided with a central processing unit (CPU)  4311 , a memory unit  4312  and a networking unit  4313  by which the CPU  4311  is communicative with the FADECs  412  and  422  and with the control system or the controller  216  of  FIG.  3   . The memory unit  4312  has executable instructions stored thereon which are readable and executable by the CPU  4311  to cause the CPU  4311  to execute the methods and operations described herein. The controller  216  of  FIG.  3    can be configured similarly. 
     With the configurations described above, during descent of the aircraft  402  and an initialization of a fuel conservation mode, the control system or the controller  216  of  FIG.  3    (or the SPU controller  431  or any other control element) can execute the operation of the first hybrid-electric engine  410  in the normal or minimum descent idle level and the operation of the second hybrid-electric engine  420  in the low-power or below minimum descent idle level or the turning of the second hybrid-electric engine  420  off as well as the simultaneous diversion of the electrical power from at least the SPU  430  (and optionally electrical power from the additional or alternative power sources  440 ) to the second electric motor  421  of the second hybrid-electric engine  420  to power the second electric motor  421  to continue to drive certain non-thermal, rotational operations of the second hybrid-electric engine  420 . This can be done by at least closing electrical contacts that are electrically interposed between the SPU  430  and the second electric motor  421 . In this way, as shown in  FIG.  3   , the electrical power generated by the SPU  430  can be drawn from the SPU  430  and directed through the rectifier  331 , through the DC power distribution bus  430 , through the bi-directional power electronics  323  and  327  and through the AC distribution busses  324  and  328  to either or both of the low-spool or high-spool electric machines  30   b  or  32   b . This diverted electrical power can thus be used to power either or both of the low-spool or high-spool electric machines  30   b  or  32   b  which can then continue to drive operations of the second gas turbine engine  20   b.    
     With reference to  FIG.  6   , a method of operating a hybrid-electric aircraft system as described above and that includes first and second hybrid-electric engines, each of which comprises an electric motor, and a supplemental power unit (SPU) configured to generate electrical power is provided. As shown in  FIG.  6   , the method includes initiating a fuel conservation mode ( 601 ), running the first hybrid-electric engine normally with the second hybrid-electric engine being turned off ( 602 ) and diverting electrical power generated by the SPU to the electric motor of the second hybrid-electric engine ( 603 ). The running of the first hybrid-electric engine normally or at the minimum descent idle level with the second hybrid-electric engine being turned off of operation  602  can be executed simultaneously with the diverting of the electrical power generated by the SPU to the electric motor of the second hybrid-electric engine during descent of operation  603 . 
     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.