Patent Publication Number: US-2023138133-A1

Title: Aircraft electric motor with integrated cooling system

Description:
STATEMENT OF FEDERAL SUPPORT 
     This invention was made with government support under Contract No. DE-AR0001351 awarded by the U.S. Department of Energy. The government has certain rights in the invention. 
    
    
     BACKGROUND 
     The present disclosure relates to electric motors, and more particularly, to electric motor assemblies with high efficiency and power density with a light weight for aircraft applications. 
     Traditional electric motors may include a stator and a rotor, with electrical motor windings in the stator that, when energized, drive rotation of the rotor about a central axis. Heat is generated in the motor windings, which are located in slots in the stator. The windings are separated from the exterior of the motor by layers of insulation and laminated steel, which makes up the stator. These contributors to internal thermal resistance limit the allowable heat generation and thus the allowable electrical current in the windings. The energy density of an electric motor is typically limited by heat dissipation from the motor windings of the stator. The requirement to be met is a maximum hot spot temperature in the motor windings that is not to be exceeded. Conventional motor thermal management includes natural convection from large fins on the outside of a motor jacket, or liquid cooling in the motor jacket. Both of these solutions undesirably add volume and/or weight to the motor, due to the addition of, at least, the jacket. 
     BRIEF DESCRIPTION 
     According to some embodiments of the present disclosure, aircraft electric motors are provided. The aircraft electric motors include a motor unit having a rotor and a stator, wherein the stator includes a plurality of windings and cooling channels arranged to provide cooling to the plurality of windings, a drive unit configured to drive operation of the motor unit, and a cooling system. The cooling system includes a working fluid arranged within a cooling fluid flow path, wherein the cooling fluid flow path includes a liquid cooling path configured to direct flow of the working fluid through, at least, the cooling channels of the motor unit and a vapor cooling path configured to direct flow of the working fluid through the drive unit and a separator arranged upstream of each of the liquid cooling path and the vapor cooling path and configured to direct a liquid portion of the working fluid into the liquid cooling path and configured to direct a vapor portion of the working fluid into the vapor cooling path. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a mixer arranged along the cooling fluid flow path and configured to merge the flow of the liquid cooling path and the vapor cooling path at a location downstream of the motor unit and the drive unit. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a heat exchanger arranged along the cooling fluid flow path and configured to cool the working fluid to a liquid state. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a mixer arranged along the cooling fluid flow path downstream of the motor unit and the drive unit and upstream of the heat exchanger, the mixer configured to merge the flow of the liquid cooling path and the vapor cooling path. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the working fluid is a first fluid of the heat exchanger and air is a second fluid of the heat exchanger, wherein the air passed through the heat exchanger to cool the working fluid. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a reservoir configured to contain the working fluid in a liquid state to be supplied to the separator. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a pump arranged upstream of the separator and configured to provide a motive force to the working fluid to flow through the cooling fluid flow path. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a control valve arranged along the liquid cooling flow path to control a flow rate of the working fluid through the liquid cooling flow path. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include a control valve arranged along the vapor cooling flow path to control a flow rate of the working fluid through the vapor cooling flow path. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the vapor cooling path includes both a vapor state of the working fluid and liquid droplets of the working fluid. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the working fluid is a saturated refrigerant. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the refrigerant is a hydrofluorocarbon (HFC), a hydrofluro-olefin (HFO), or a hydrofluoroether (HFE). 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the windings are arranged in a U-shape configuration. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the motor unit comprises rotor having U-shaped magnets arranged about the windings of the stator. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the separator is part of a header fluidly connected to the cooling channels. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that at least one cooling channel includes a restrictor configured to control flow through the respective cooling channel. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the liquid portion is at least 75% liquid state working fluid. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the vapor portion is at least 95% vapor state working fluid. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the rotor and stator are arranged in an annular configuration. 
     In addition to one or more of the features described herein, or as an alternative, further embodiments of the aircraft electric motors may include that the stator is arranged within the rotor. 
     The foregoing features and elements may be executed or utilized in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Various features will become apparent to those skilled in the art from the following detailed description of the disclosed non-limiting embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
         FIG.  1 A  is a partial view of an embodiment of electric motor; 
         FIG.  1 B  is a cross-sectional view of an embodiment of a stator core of the electric motor of  FIG.  1 A ; 
         FIG.  2 A  is a schematic illustration of an aircraft electric motor in accordance with an embodiment of the present disclosure; 
         FIG.  2 B  is a side elevation view of the aircraft electric motor of  FIG.  2 A ; 
         FIG.  2 C  is a partial cut-away illustration of the aircraft electric motor of  FIG.  2 A ; 
         FIG.  2 D  is a separated-component illustration of the aircraft electric motor of  FIG.  2 A ; 
         FIG.  3 A  is a schematic illustration of a rotor and stator of an aircraft electric motor in accordance with an embodiment of the present disclosure; 
         FIG.  3 B  is a schematic illustration of the rotor and stator of  FIG.  3 A  as arranged within a rotor sleeve in accordance with an embodiment of the present disclosure; 
         FIG.  4    is a schematic illustration of an aircraft electric motor system in accordance with an embodiment of the present disclosure; 
         FIG.  5    is a schematic view of a power system of an aircraft that may employ embodiments of the present disclosure. 
     
    
    
     The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, that the following description and drawings are intended to be illustrative and explanatory in nature and non-limiting. 
     DETAILED DESCRIPTION 
     Referring to  FIGS.  1 A- 1 B , schematic illustrations of an electric motor  100  that may incorporate embodiments of the present disclosure are shown.  FIG.  1 A  illustrates a cross-sectional view of the electric motor  100  and  FIG.  1 B  illustrates a cross-sectional view of a stator core of the electric motor  100 . The electric motor  100  includes a rotor  102  configured to rotate about a rotation axis  104 . A stator  106  is located radially outboard of the rotor  102  relative to the rotation axis  104 , with a radial air gap  108  located between the rotor  102  and the stator  106 . As illustrated, the rotor  102  may be mounted on a shaft  110  which may impart rotational movement to the rotor  102  or may be driven by rotation of the rotor  102 , as will be appreciated by those of skill in the art. The rotor  102  and the shaft  110  may be fixed together such that the rotor  102  and the shaft  110  rotate about the rotation axis  104  together as one piece. 
     The stator  106  includes a stator core  112  in which a plurality of electrically conductive stator windings  114  are disposed. In some embodiments, such as shown in  FIG.  1 A , the stator core  112  is formed from a plurality of axially stacked laminations  116 , which are stacked along the rotation axis  104 . In some embodiments, the laminations  116  are formed from a steel material, but one skilled in the art will readily appreciate that other materials may be utilized. The stator windings  114 , as shown, include core segments  118  extending through the stator core  112  and end turn segments  120  extending from each axial stator end  122  of the stator core  112  and connecting circumferentially adjacent core segments  118 . When the stator windings  114  are energized via an electrical current therethrough, the resulting field drives rotation of the rotor  102  about the rotation axis  104 . Although  FIG.  1 A  illustrates the stator core  112  arranged radially inward from the stator windings  114 , it will be appreciated that other configurations are possible without departing from the scope of the present disclosure. For example, in some embodiments, the stator structure may be arranged radially inward from a rotating rotor structure. In some embodiments, the rotor may be both radially inward and outward of the stator, as described herein. 
       FIG.  1 B  is an axial cross-sectional view of the stator core  112 . Each lamination  116  of the stator core  112  includes a radially outer rim  124  with a plurality of stator teeth  126  extending radially inwardly from the outer rim  124  toward the rotation axis  104 . Each of the stator teeth  126  terminate at a tooth tip  128 , which, together with a rotor outer surface  130  (shown in  FIG.  1 A ) of the rotor  102 , may define the radial air gap  108 . Circumferentially adjacent stator teeth  126  define an axially-extending tooth gap  132  therebetween. Further, in some embodiments, a plurality of stator fins  134  extend radially outwardly from the outer rim  124 . 
     Electric motors, as shown in  FIGS.  1 A- 1 B  may require cooling due to high density configurations, various operational parameters, or for other reasons. For example, high-power-density aviation-class electric motors and drives may require advanced cooling technologies to ensure proper operation of the motors/drives. These machines are generally thermally limited at high power ratings and their performance can be improved by mitigating thermal limitations. To maintain desired temperatures, a thermal management system (TMS) is integrated into the system, which provides cooling to components of the system. 
     Onboard an aircraft, power requirements, and thus thermal management system (TMS) loads, are substantially higher during takeoff. Sizing of the TMS for takeoff conditions (i.e., maximum loads) results in a TMS having a high weight to accommodate such loads. This results in greater weight and lower power density during cruise conditions which do not generate such loads, and thus does not require a high cooling capacity TMS. Balancing weight constraints and thermal load capacities is important for such aviation applications. 
     In view of such considerations, improved aviation electric motors are provided herein. The aviation electric motors or aircraft electric motors, described herein, incorporate lightweight materials and compact design to reduce weight, improve thermal efficiencies, improve power efficiencies, and improve power density. 
     Turning now to  FIGS.  2 A- 2 D , schematic illustrations of an aircraft electric motor  200  in accordance with an embodiment of the present disclosure are shown.  FIG.  2 A  is an isometric illustration of the aircraft electric motor  200 ,  FIG.  2 B  is a side elevation view of the aircraft electric motor  200 ,  FIG.  2 C  is a partial cut-away view illustrating internal components of the aircraft electric motor  200 , and  FIG.  2 D  is a schematic illustration of components of the aircraft electric motor  200  as separated from each other. The aircraft electric motor  200  includes a motor housing  202 , a cooling system  204 , a first power module system  206 , and a second power module system  208 . 
     The motor housing  202  houses a stator  210  and a rotor  212 , with the rotor  212  configured to be rotatable about the stator  210 . In this illustrative embodiment, the rotor  212  includes a U-shaped magnet  214  arranged within a similarly shaped U-shaped rotor sleeve  216 . The rotor sleeve  216  is operably connected to a hub  218 . The hub  218  is fixedly attached to a first shaft  220 . The first shaft  220  is operably connected to a second shaft  222 . In some configurations, the first shaft  220  may be a high speed shaft and may be referred to as an input shaft. In such configurations, the second shaft  222  may be a low speed shaft and may be referred to as an output shaft. The connection between the first shaft  220  and the second shaft  222  may be by a gear assembly  224 , as described herein. 
     The cooling system  204  is configured to provide cooling to the components of the aircraft electric motor  200 . The cooling system  204 , as shown in  FIG.  2 D , includes a heat exchanger  226  and a header  228 . The heat exchanger  226  and the header  228  may form a closed-loop cooling system that may provide air-cooling to a working fluid at the heat exchanger  226 . The header  228  may be, in some configurations, a two-phase di-electric cooling header. A cooled working fluid may be pumped from the heat exchanger  226  into the header  228  using a pump  229  and distributed into embedded cooling channels  230  that are arranged within the stator  210 . As the aircraft electric motor  200  is operated, heat is generated and picked up by the working fluid within the embedded cooling channels  230 . This heated working fluid is then passed through the header  228  back to the heat exchanger  226  to be cooled, such as by air cooling. Although described as air-cooling, other cooling processes may be employed without departing from the scope of the present disclosure. 
     As shown, the heat exchanger  226  of the cooling system  204  may be a circular structure that is arranged about the motor housing  202 . This configuration and arrangement allows for improved compactness of the system, which may be advantageous for aircraft applications. The rotor sleeve  216  with the magnets  214 , the stator  210 , and the gear assembly  224  fit together (although moveable relative to each other) within the motor housing  202 , providing for a compact (low volume/size) design. 
     As noted above, the rotor sleeve  216  may be operably coupled to a first shaft  220  by the hub  218 . The first shaft  220  may be operably coupled to a first gear element  232  and the second shaft  222  may be operably coupled to a second gear element  234 . The first and second gear elements  232 ,  234  may form the gear assembly  224 . The first and second gear elements  232 ,  234  are arranged to transfer rotational movement from the first shaft  220 , which is driven in rotation by the hub  218  and the rotor sleeve  216  of the rotor  212 , to the second shaft  222 . In some embodiments, the first shaft  220  may be operably connected to a sun gear as the first gear element  232  that engages with a plurality of planetary gears and drives rotation of the second gear element  234  which may be operably connected to the second shaft  222 . In some embodiments, the second shaft  222  may be connected to a fan or other component to be rotated by the aircraft electric motor  200 . 
     The aircraft electric motor  200  includes the first power module system  206  and the second power module system  208 . The first and second power module systems  206 ,  208  can include capacitors and other electronics, including, but not limited to, printed circuit boards (PCBs) that may enable control and operation of the aircraft electric motor  200 . Again, the profile of the aircraft electric motor  200  of the present disclosure presents a low profile or compact arrangement that reduces the volume of the entire power system, which in turn can provide for improved weight reductions. In some embodiments, the first and second power module systems  206 ,  208  may be electrically connected to the stator  210  to cause an electric current therein. As the electric current will induce an electromagnetic field which will cause the rotor  212  to rotate. 
     Referring now to  FIGS.  3 A- 3 B , schematic illustrations of a portion of an aircraft electric motor  300  in accordance with an embodiment of the present disclosure is shown.  FIGS.  3 A- 3 B  illustrate a portion of a rotor  302  and a stator  304  of the aircraft electric motor  300 .  FIG.  3 A  illustrates the rotor  302  and the stator  304  and  FIG.  3 B  illustrates these components arranged within a rotor sleeve  306 . 
     The rotor  302  is formed of a plurality of U-shaped magnets  308 . In some configurations, the plurality of magnets  308  can be arranged with alternating polarity in a circular or annular structure. Arranged within the “U” of the U-shaped magnets  308  is the stator  304 . The stator  304  is formed of a plurality of windings  310 . In this configuration, the windings  310  are arranged with a header  312 . The header  312  may be part of a cooling system, such as that shown and described above. The header  312  can be configured to cycle a working fluid through cooling channels  314  for cooling of the windings  310 , as shown in  FIG.  3 B . As shown in  FIG.  3 B , the cooling channels  314  may include a flow restrictor  315  arranged at an inlet side (or an outlet side) of the cooling channel  314 . The flow restrictor  315  may be used to throttle the flow of a cooling fluid to provide efficient cooling within the cooling channels  314 . The flow restrictor  315  may be configured to manage the location of subcooled liquid and/or flow boiling within the cooling channels  314 . 
     The windings  310  may be wrapped about a support structure  316 . The support structure  316 , in some embodiments and as shown in  FIG.  3 B , may include a laminate portion  318  and a magnetic portion  320 . In some such embodiments, the laminate portion  318  may be formed from cobalt steel laminate and the magnetic portion  320  may be formed from a soft magnetic composite. The laminate portion  318  may be provided to capture in-plane flux from outer and inner rotor. The magnetic portion  320  may be provided to capture end rotor flux and may take a shape/filler in a gap through the end turns of the coil. The windings  308  include end connections  322  and may be electrically connected to one or more power module systems of the aircraft electric motor, such as shown above. 
     As shown in  FIG.  3 B , the magnets  306  are U-shaped and arranged within the rotor sleeve  306 . The rotor sleeve  306  is a substantially U-shaped sleeve that is sized and shaped to receive the U-shaped magnets  308 . In this illustrative configuration, the rotor sleeve  306  can include an inner sleeve  324 . The inner sleeve  324  may be configured to provide support to a portion of the magnets  308 . It will be appreciated that there is no direct contact between the windings  310  and the magnets  308 . This lack of contact enables free rotation of the rotor  302  relative to the stator  304  during operation. 
     High-power-density aviation-class electric motor and drives, such as those shown and described above, may require advanced cooling technologies. These machines are generally thermally limited at high power ratings and their performance can be improved by mitigating thermal limitations. Accordingly, embodiments of the present disclosure are directed to improved cooling schemes for aircraft electric motors (e.g., as described above). Embodiments of the present disclosure are directed to employing a two-phase cooling scheme to improve cooling at high load locations (e.g., within windings of the motor). Two-phase cooling is a highly efficient approach for cooling the heat generating components. Non-uniform flow (e.g., liquid/vapor phase) distribution, where some channels receive insufficient liquid coolant, is a critical risk in a two-phase cooling approach. Both to improve flow distribution in motor channels and to optimize overall performance of the thermal management system, the loop architecture in accordance with embodiments of the present disclosure can be optimized so that the more critical components receive more coolant liquid (as compared to vapor or a mixture). 
     Turning now to  FIG.  4   , a schematic illustration of a motor system  400  in accordance with an embodiment of the present disclosure is shown. The motor system  400  may include an aircraft electric motor  402  that is operably connected to a cooling system, as described herein. The aircraft electric motor  402  includes a motor unit  402  and a drive unit  404 . The motor unit  404  may include windings, magnets, support structures, etc. as shown and described above. The drive unit  406  may include power modules, electronic components, and the like. 
     To provide cooling to the motor unit  404  and/or the drive unit  406 , the motor system  400  includes a cooling scheme that has a cooling fluid flow path  408 . A working fluid may be passed through the cooling fluid flow path  408  to provide cooling to components arranged along the cooling fluid flow path  408  (e.g., motor unit  404 , drive unit  406 , etc.). The working fluid may be a saturated refrigerant (e.g., dielectric refrigerants including, but not limited to, hydrofluorocarbons (HFC), hydrofluro-olefins (HFO), and/or hydrofluoroethers (HFE)). As the working fluid passes through the cooling fluid flow path  408 , the working fluid may undergo phase change such that both vapor and liquid components of the working fluid may pass through portions of the cooling fluid flow path  408 . Because of the increased load capacity of liquid (as compared to vapor phases), it may be preferable to have a substantially pure liquid phase of the working fluid pass through and cool the motor unit  404 . To achieve such substantially pure liquid phase of the working fluid, the motor system  400  includes a separator  410  arranged upstream of the motor unit  404  and the drive unit  406 . 
     The separator  410  is configured to separate the working fluid into two flow streams along a liquid cooling path  412  and a vapor cooling path  414  (of the working fluid flow path  408 ). The separator  410  is configured to create a liquid supply  416  of the working fluid to be conveyed along the liquid cooling path  412  and a vapor supply  418  of the working fluid to be conveyed along the vapor cooling path  414 . It will be appreciated that the liquid supply  416  may include some entrained or portion that is still in vapor form, but such percentage of volume is minimal. For example, the liquid supply  416  may be configured to supply a working fluid being at least 75% (by volume) liquid working fluid to the motor unit  404 . Similarly, the vapor supply  418  may include some entrained or portion that is still in liquid form (e.g., liquid droplets), but such percentage of volume is minimal. For example, the vapor supply  416  may be configured to supply a working fluid being at least 95% (by volume) vapor working fluid to the drive unit  404 . 
     After providing cooling to the aircraft electric motor  402 , the two flows of the working fluid may be rejoined at a mixer  420 . The mixer  420  may be a manifold or other structure where the flow along the liquid cooling path  412  and the vapor cooling path  414  are rejoined or combined to create a single flow of working fluid downstream from the aircraft electric motor  402 . The flow through the liquid cooling path  412  may be controlled by an optional control valve  422  arranged along the liquid flow path  412  and the flow through the vapor cooling path  414  may be controlled by an optional control valve  424  arranged along the vapor flow path  414 . 
     The merged or joined flows of the working fluid, at the mixer  420 , are then passed into and through a heat exchanger  424  (e.g., a condenser). The heat exchanger  424  receives the working fluid of the motor system  400  as a first fluid and a cooling fluid  426  as a second fluid. The cooling fluid  426  may be air or other fluid, such as a refrigerant, a gas, a liquid, etc. In some embodiments, the heat exchanger  424  may be substantially similar to the heat exchanger  226  with the header  228  shown and described above. In other embodiments, the heat exchanger  424  may be arranged separate from the motor system  400  (e.g., not arranged about the motor system). As the working fluid of the motor system  400  passes through the heat exchanger  424 , the temperature of the working fluid will decrease, thus causing most or all the working fluid to either remain or phase change to a liquid state. The cooled (and liquid) working fluid may then be passed into a reservoir  428 . The reservoir  428  may be configured to collect and store liquid state working fluid of the motor system  400 . 
     A pump  430  is arranged along the working fluid flow path  408 . The pump  430  is configured to impart a motive force to cause the working fluid to flow along the working fluid flow path  408 . In some embodiments, the pump  430  may be configured to extract liquid working fluid from the reservoir  428  to drive the working fluid through the aircraft electric motor  402  and provide cooling thereto. Although the pump  430  is configured to pump a liquid state fluid, the pumping action and the nature of a saturated refrigerant may cause vapor to form within the working fluid flow path  408 . Thus, the separator  410  is arranged to separate these two states of the working fluid and direct them along the respective liquid cooling path  412  and vapor cooling path  414 . 
     As a result of the disclosed working fluid flow path  408  using a two-phase refrigerant, liquid (higher load) working fluid (e.g., cooling fluid) may be directed to the components that require the highest heat removal (e.g., motor unit  404 ). That is, by employing a split-line, two-phase cooling scheme, embodiments of the present disclosure can ensure more critical components receive more coolant liquid (as compared to vapor). 
     In operation of the motor system  400 , a saturated refrigerant is separated into vapor and liquid phases at the separator  410 . The liquid portion is fed to motor winding cooling channels within the motor unit  404  along the liquid cooling path  412 . The vapor portion, and any remaining liquid droplets, are used to cool the less demanding drive electronics. Such as the drive unit  406  along the vapor cooling path  414 . Subcooled liquid flow entering cooling channels of the motor unit  404  may be throttled through a restriction at the inlet of the channels in order to provide efficient cooling thereto. That is, the liquid cooling flow path  412  may include one or more flow restrictors at the inlet side of cooling channels within windings of a motor (e.g., at an inlet of cooling channels  314  for cooling of the windings  310 ). The header manifold (e.g., header  228 ) feeding motor channels with the working fluid may act as the separator (e.g., separate  410 ) where vapor is extracted at the top and fed to components of the drive unit  406 . In some embodiments, through control of the control valves  420 ,  422 , an actively controlled flow of the working may be achieved. Such control can ensure that more liquid is directed to hot spots as needed. For example, if excess heat is detected or if some components are overloaded in case of failure of other components, then additional or preferential flow of cooling may be directed into and through the liquid cooling flow path  412  and into the motor unit  404 . 
     Referring now to  FIG.  5   , a power system  500  of an aircraft  502  is shown. The power system  500  includes one or more engines  504 , one or more electric motors  506 , a power bus electrically connecting the various power sources  504 ,  506 , and a plurality of electrical devices  510  that may be powered by the engines  504  and/or motors  506 . The power system  500  includes a power distribution system  512  that distributes power  514  through power lines or cables  516 . The electric motors  506  of the aircraft  502  may be configured similar to the aircraft electric motors shown and described above. 
     Advantageously, embodiments of the present disclosure provide for improved electric motors for aircraft and aviation applications. The aircraft electric motors of the present disclosure have improved cooling, which may enable higher operating temperatures and/or loads. Further, because of the separation of the liquid and vapor phases of the cooling working fluid, improved flow distribution through the motor cooling channels may be achieved. The systems described herein may be closed loop, even with the inclusion of the separation of the two phases of the working fluid, and through one or more controller and/or valves, preferential cooling may be achieved by directing liquid cooling fluid to components that may require an increased cooling load (e.g., excessive heat, compensation for a failed component, etc.). 
     The terms “about” and “substantially” are 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. For example, “about” or “substantially” can include a range of ±8% or 5%, or 2% of a given value. 
     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.