Patent Publication Number: US-2023141526-A1

Title: Aircraft electric motor

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 having at least one directional feature forming a portion of at least one cooling channel, the at least one directional feature configured to prevent backflow of a cooling fluid that passes through the at least one 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 at least one directional feature comprises a plurality of directional features distributed between an inlet of the at least one cooling channel to an outlet of the at least one 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 at least one directional feature is a Tesla valve. 
     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 at least one directional feature is a series of helical grooves. 
     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 at least one directional feature is a series of rings, wherein each ring has an angled sidewall that has a larger diameter than an upstream directional feature at the upstream end of the angled sidewall. 
     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 at least one cooling channel defines an inlet portion, a mid-channel portion, and an outlet portion. 
     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 at least one directional feature is located within the inlet portion. 
     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 at least one directional feature is located within the outlet portion. 
     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 at least one directional feature is located within the mid-channel portion. 
     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 at least one directional feature comprises at least one feature in the inlet portion and at least one feature in the outlet portion. 
     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 mid-channel portion does not include any directions features. 
     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 at least one feature within the inlet portion is different from the at least one feature in the outlet portion. 
     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 at least one feature within the inlet portion is the same as the at least one feature in the outlet portion. 
     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 cooling system further includes a header and a heat exchanger configured to supply cooling fluid into the plurality of 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 a pump configured to pump the cooling fluid from the header into the plurality of 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 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 at least one directional feature is configured to cause a phase change in the cooling 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 cooling fluid is one of a hydrofluorocarbon (HFC), a hydrofluro-olefin (HFO), or a hydrofluoroether (HFE). 
     According to some embodiments, 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 means for driving operation of the motor unit, and a cooling system. The cooling system includes at least one means for supplying cooling fluid to a portion of the stator and at least one means for preventing backflow of the cooling fluid within the at least one means for supplying cooling 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 means for driving operation of the motor unit comprises at least one power module system, the at least one means for supplying cooling fluid comprises at least one cooling channel, and the at least one means for preventing backflow is at least one directional feature formed in a respective one of the at least one cooling channel. 
     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 a portion of an aircraft electric motor system in accordance with an embodiment of the present disclosure; 
         FIG.  5    is a schematic illustration of a portion of an aircraft electric motor system in accordance with an embodiment of the present disclosure; 
         FIG.  6    is a schematic illustration of a portion of an aircraft electric motor system in accordance with an embodiment of the present disclosure; 
         FIG.  7    is a schematic illustration of a portion of an aircraft electric motor system in accordance with an embodiment of the present disclosure; and 
         FIG.  8    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. 
       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 . As such, the first and second power module systems  206 ,  208  may form at least a part of a drive unit of the aircraft electric motor  200 . 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 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). 
     High-power-density aviation-class electric motor and drives, as 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. Two-phase cooling is a highly efficient approach for cooling the heat generating components. In accordance with embodiments of the present disclosure, a cooling refrigerant is configured to boil and evaporate within embedded micro-channels in the electric components such as the winding. The boiling of the refrigerant within he microchannels can result in reverse flow (backflow) in channels and subsequent local dry out. As such, embodiments of the present disclosure are configured to ensure a consistent flow direction and minimize or prevent backflow of the fluid. 
     In accordance with embodiments of the present disclosure, channel surfaces of cooling channels embedded or formed within windings of an aircraft electric motor are enhanced with three dimensional features that both enhance heat transfer and preferentially direct flow in one direction. These features may have much higher flow resistance (pressure drop) in one direction compared to the reverse direction. The surfaces/channels described herein may be fabricated through additive processes, machining, casting, or through other manufacturing processes. The features described herein may be local. For example, such features may be arranged near an entrance/inlet of a channel or distributed intermittently along the channel, etc. In other configurations, the features may be present along a full length from inlet to outlet of a given cooling channel. The features may be sized locally based on vapor quality. 
     Turning now to  FIG.  4   , a schematic illustration of example flow channels for use in an aircraft electric motor  400  in accordance with an embodiment of the present disclosure is shown.  FIG.  4    illustrates a portion of a first winding  402  and a second winding  404  that may be part of a stator portion of an aircraft electric motor of the present disclosure. The first winding  402  includes a respective first preferential direction cooling channel  406  and the second winding  404  includes a respective second preferential direction cooling channel  408 . As shown, each of the first preferential direction cooling channel  406  and the second preferential direction cooling channel  408  are configured as Tesla valve configurations. 
     In  FIG.  4   , the first preferential direction cooling channel  406  has a first flow configuration  410  which has a cooling fluid  412  flow from an inlet  414  to an outlet  416 . In the first flow configuration  410 , the fluid flow of the cooling fluid  412  is in a preferential flow direction (from inlet  414  to outlet  416 ) that minimizes trapping fluid or causing backflow. The first preferential direction cooling channel  406  includes branches  418  which may allow for pockets of boiling cooling fluid  412  to expand without causing the turbulent flow that would induce or cause backflow. The cooling fluid may be a saturated refrigerant (e.g., dielectric refrigerants including, but not limited to, hydrofluorocarbons (HFC), hydrofluro-olefins (HFO), and/or hydrofluoroethers (HFE)). 
     In contrast, in a second flow configuration  420 , the flow of a cooling fluid  422  is from an inlet  424  toward an outlet  426  that is the opposite of that in the first flow configuration  410 . The geometric shape of the second preferential direction cooling channel  408  is the same as that of the first preferential direction cooling channel  406 , but the flow direction is reversed. As such, when a portion  428  of the cooling fluid  422  enters branches  430  (e.g., due to boiling of the cooling fluid  422 ), the fluid may flow about and reverse direction and cause a turbulent flow at a junction  432 . This turbulence is the result of a backflow or a flow of the cooling fluid  422   flowing in a direction opposite of the primary flow (e.g., in a direction from the outlet  426  toward the inlet  424 ). 
     The first flow configuration  410  is a preferential flow, such that the orientation of the features of the first preferential direction cooling channel  406  are arranged to minimize backflow while allowing for improved cooling to the surrounding first winding  402 . By ensuring a preferential flow direction and minimal disruption, improved heat removal may be achieved, thus increasing the amount of cooling provided by the cooling channels. In the configuration of  FIG.  4   , the cooling channels are arranged as Tesla valves. Such valves allow fluid to flow unimpeded in one direction (e.g., inlet  414  toward outlet  416 ), but in the other direction, the fluid is blocked or prevented from flowing in such opposite direction. The interior of the cooling channels are provided with enlargements, recesses, projections, baffles, or buckets which, while offering virtually no resistance to the passage of the cooling fluid in one direction, other than surface friction, constitute an almost impassable barrier to its flow in the opposite direction. 
     The preferential direction cooling channels of the present disclosure are not limited to Tesla valve configurations. For example, toothed configurations, groves, helical grooves or embedded channels, rings, or the like may be used to provide a preferential direction feature within and/or along a cooling channel. 
     For example, with reference to  FIG.  5   , a schematic illustration of a portion of an aircraft electric motor  500  in accordance with an embodiment of the present disclosure is shown. In  FIG.  5   , a winding  502  is shown having a preferential flow direction cooling channel  504  passing therethrough. The winding  502  may be part of a stator of an aircraft electric motor of the present disclosure, such as shown and described above. The preferential flow direction cooling channel  504  defines a preferential flow direction Flow p . The preferential flow direction cooling channel  504  is defined by directional features  506  that include, in this embodiment, a flow prevention element  508 . The flow prevention element  508  is a structural feature of the exterior wall defining the preferential flow direction cooling channel  504 . The flow prevention element  508  of a given directional feature  506  may be a land of a thread or ring shape that has a larger diameter than an upstream directional feature  506 . As such, flow in a direction counter to the preferential flow direction Flow p  will be discouraged (e.g., minimized or eliminated). In the ring configuration, each ring structure has an angled sidewall that has a larger diameter than an upstream directional feature at the upstream end of the angled sidewall. 
     With reference to  FIG.  6   , a schematic illustration of a portion of an aircraft electric motor  600  in accordance with an embodiment of the present disclosure is shown. In  FIG.  6   , a winding  602  is shown having a preferential flow direction cooling channel  604  passing therethrough. The winding  602  may be part of a stator of an aircraft electric motor of the present disclosure, such as shown and described above. The preferential flow direction cooling channel  604  defines a preferential flow direction Flow p . The preferential flow direction cooling channel  604  is defined by directional features  606  that include, in this embodiment, a flow prevention element  608 . The flow prevention element  608  is a structural feature of the exterior wall defining the preferential flow direction cooling channel  604 . The flow prevention element  608  of a given directional feature  606  may be a flat face of a tooth or thread that has a smaller diameter than an upstream directional feature  606 . As such, flow in a direction counter to the preferential flow direction Flow p  will be discouraged (e.g., minimized or eliminated). 
     With reference to  FIG.  7   , a schematic illustration of a portion of an aircraft electric motor  700  in accordance with an embodiment of the present disclosure is shown. In  FIG.  7   , a winding  702  is shown having a preferential flow direction cooling channel  704  passing therethrough. The winding  702  may be part of a stator of an aircraft electric motor of the present disclosure, such as shown and described above. The preferential flow direction cooling channel  704  defines a preferential flow direction Flow p . The preferential flow direction cooling channel  704  is defined by directional features  706 ,  708 . A first set of directional features  706  are arranged proximate an inlet end  710  of the preferential flow direction cooling channel  704  and a second set of directional features  708  are arranged proximate an outlet end  712  of the preferential flow direction cooling channel  704 . The preferential flow direction cooling channel  704  also includes a portion that has not preferential direction features. As such, in this illustratively embodiment, the preferential flow direction cooling channel  704  has an inlet portion  714  including the first set of directional features  706 , a mid-channel portion  716  having no directional features, and an outlet portion  718  including the second set of directional features  708 . 
     In this illustrative embodiment, the first set of directional features  706  have a different structural configuration than the second set of directional features  708 . In other embodiments, the directional features at the inlet end  710  and the outlet end  712  may be substantially similar. The selection of the specific configuration of the directional features  706 ,  708  at the inlet portion  714 , the outlet portion  718 , or the inclusion or absence of such features in the mid-channel portion  716 , may be selected to achieve a desired cooling flow and cooling load. Further, the Tesla valve configuration illustrated in  FIG.  4    may be combined with other (different) directional features, as will be appreciated by those of skill in the art in view of the teachings herein. Further, the length of the respective portions  714 ,  716 ,  718  may be varied to achieve desired cooling properties. For example, a series of different (or similar) directional features may be distributed along the length of a cooling channel, with sections or portions of smooth channel between the sets of directional features. 
     Referring now to  FIG.  8   , a power system  800  of an aircraft  802  is shown. The power system  800  includes one or more engines  804 , one or more electric motors  806 , a power bus electrically connecting the various power sources  804 ,  806 , and a plurality of electrical devices  810  that may be powered by the engines  804  and/or motors  806 . The power system  800  includes a power distribution system  812  that distributes power  814  through power lines or cables  816 . The electric motors  806  of the aircraft  802  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 channels that may improve cooling while eliminating or at least reducing the challenges with cooling windings of aircraft electric motors. For example, advantageously, embodiments of the present disclosure may prevent backflow within the cooling fluid flow, which may be detrimental or at least reduce efficiencies of the cooling. 
     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.