Patent Publication Number: US-2016229525-A1

Title: Electromechanical rotary actuator

Description:
BACKGROUND OF INVENTION 
     This invention relates, generally, to an actuator and, more specifically, to an electromechanical hinge-line rotary actuator for use with a thin-wing aircraft in flight-control applications. 
     Many systems require actuators to manipulate various components. Rotary actuators rotate an element about an axis. In flight-control applications, there has been a trend toward a thinner wing such that size and space are limited at a point of attachment between the wing and an aileron (a wing-control surface) of an aircraft. 
     This trend has driven use of a rotary actuator of a “hinge-line” design, wherein a rotational axis of the actuator is aligned with that of the aileron and the actuator acts as a hinge (hence, the term “hinge-line”). This trend also raises a need for such an actuator with a tighter cross-section, which limits the diameter of a motor of the actuator, and higher power density. 
     In turn, torque of the motor is directly related to the motor diameter and current flowing through windings of the motor. However, with the limited motor diameter and an amount of the current being limited to useable amounts on a power bus of the aircraft, an amount of such torque is limited as well. And, since power of the motor equates to speed thereof times the torque amount and this amount is limited, the speed must be higher. Yet, use of the higher-speed motor at the limited torque amount is driving use of higher gear ratios, which makes inertia of the motor a sensitive design parameter. 
     More specifically, reflected inertia comes into play whenever the motor or a gear set of the aircraft is trying to be back-driven, which is a requirement for a surface of the aileron. And, reduction in the inertia prior to a gear affects the reflected inertia by a factor of a gear ratio squared (for example, a “10:1” gear ratio yields a reflected inertia of 100 times greater than the motor inertia while a “100:1” gear ratio yields a reflected inertia of 10,000 times greater). The inertia also affects responsiveness of the aircraft—i.e., a higher level of the inertia results in a lower responsiveness. 
     A typical electromechanical hinge-line rotary actuator designed for flight-control applications is arranged to use a conventional motor that is framed (i.e., encased, housed, or mounted) and includes a rotor. The rotor is disposed inside the frame and indirectly connected to an end of a planetary gearbox or gear set through a drive shaft or coupler. In this way, the motor is disposed exterior to and in alignment with the gear set, and there are bearings for the motor and gear set. Such alignment is accomplished by a precision-machined housing for the motor and gear set or compliant coupling on an output shaft of the motor to an input of the gear set. This arrangement has inefficiencies associated with packaging and is not optimized for typical requirements of such an actuator. More specifically, it is not optimized for power density, performance, and reliability. 
     Accordingly, it is desirable to provide an electromechanical hinge-line rotary actuator an arrangement of which does not have inefficiencies associated with packaging and is optimized for typical requirements of such an actuator in flight-control applications. More specifically, it is desirable to provide such an actuator that reduces inertia and is optimized for power density, performance, and reliability. 
     BRIEF DESCRIPTION OF INVENTION 
     According to a non-limiting exemplary embodiment of the invention, an electromechanical rotary actuator is provided. The actuator includes a drive member and a motor disposed inside and directly coupled to the drive member. The motor has a rotor configured toward an outside of the motor and directly coupled to an input of the drive member and a stator configured toward an inside of the motor and positioned inside the rotor. The drive member, rotor, and stator are arranged concentrically with each other. 
     The actuator is configured to be employed with a thin-wing aircraft. Toward that end, arrangement of the actuator does not have inefficiencies associated with packaging and is optimized for typical requirements of such an actuator in flight-control applications—power density, performance, and reliability. More specifically, the concentric packaging of components [i.e., the drive member and motor (stator and rotor)] of the actuator provides a higher power density. Also, a load path of the actuator is a direct drive such that a drive shaft is not required, resulting in a lower inertia and, in turn, higher performance. Furthermore, the actuator has few components (including removal of one set of bearings and no requirement as well for the compliant coupling or precision-machined housing), which lends itself to higher reliability and reduced cost. In addition, a total axial stack length of the actuator can be changed to accommodate a higher output load, making the actuator versatile for different applications. Moreover, the actuator can achieve higher forces while it maintains a same cross-section thereof, making the actuator versatile for the different applications. 
    
    
     
       BRIEF DESCRIPTION OF DRAWING 
       The subject matter that is regarded as the invention is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawing in which: 
         FIG. 1  is an end view of a non-limiting exemplary embodiment of a wing of an aircraft provided with an electromechanical hinge-line rotary actuator according to the invention. 
         FIG. 2  is a schematic top view of a non-limiting exemplary embodiment of the electromechanical hinge-line rotary actuator according to the invention. 
         FIG. 3  is a schematic side environmental view of the embodiment of the electromechanical hinge-line rotary actuator illustrated in  FIG. 2 . 
         FIG. 4  is a schematic sectional side view of the embodiment of the electromechanical hinge-line rotary actuator illustrated in  FIG. 2 . 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
     Referring now to  FIG. 1 , a non-limiting exemplary embodiment of a wing of an aircraft (not shown) is generally indicated at  10 . Although the wing  10  is disclosed herein as being implemented with a non-rotary-wing aircraft, such as an airplane, it should be appreciated that the wing  10  can be implemented with any suitable type of aircraft, in general, and non-rotary-wing or rotary-wing aircraft (such as a helicopter), in particular. 
     As shown in  FIG. 1 , the wing  10  is one of two substantially similar wings of a lift system of the aircraft (in contrast, a rotor blade would be one of a plurality of substantially similar rotor blades of a rotor system of a helicopter). The wing  10  defines a root portion (not shown) that extends to tip portion (not shown) through an aileron portion, generally indicated at  14 , which acts as a flight-control or an output-control surface (such as a wing flap). The aileron portion  14  also defines, in turn, an axis of motion or rotation  16  and includes a spar, generally indicated at  18 . The wing  10  defines further first and second opposing surfaces  20 ,  22 , a trailing edge  24 , and an opposing, leading edge  26  and includes a rearward spar, generally indicated at  28 . 
     The wing  10  includes also a control system (not shown) that has an electromechanical hinge-line rotary actuator, generally indicated at  30 , and a controller (not shown). The actuator  30  defines the axis of rotation  16 . The controller may be mounted to or near the actuator  30  and is operatively linked to the actuator  30  and a control system (not shown). 
     A stationary attachment bracket or ground arm, generally indicated at  46 , of the actuator  30  is mounted to the wing rearward spar  28  and configured to be attached to interior structure of the wing  10 . A rotatable attachment bracket or an output arm, generally indicated at  48 , of the actuator  30  is mounted to a frame of or within an interior of the aileron portion  14 . The mounting is highly flexible as long as the axis of rotation  16  of the aileron portion  14  is aligned with an axis of rotation  16  of the actuator  30 . The actuator  30  allows wing flexing and, hence, does not put undue stress on the wing  10  at points of attachment when flex is encountered, such as during turbulence. 
     It should be appreciated that the control system can also define a plurality of control surfaces (not shown) arranged within the aileron portion  14  and selectively deployed between the first and second surfaces  20 ,  22  to affect flight dynamics of the wing  10 . Each surface defines first and second surface portions. The actuator  30  is configured to rotate the surface from a first or neutral position, such that the surface is disposed within the wing  10 , to a second or deployed position, such that the surface extends out an outer periphery of the wing  10 . At this point, it should be appreciated that the above description is provided for the sake of completeness and to enable a better understanding of one non-limiting exemplary application of the actuator  30 . 
     Referring now to  FIGS. 2-4 , a non-limiting exemplary embodiment of the actuator  30  is shown. The actuator  30  is disclosed herein as being implemented with a control system for a flight-control application. However, it should be appreciated that the actuator  30  can be implemented in any suitable system capable of operating in multiple environments and should not be considered as being limited to non-rotary or rotary aircraft or aircraft of any kind. 
     The actuator  30  includes, in general, a drive member, generally indicated at  36 , a motor, generally indicated at  38  ( FIG. 1 ), disposed inside and directly coupled to the drive member  36 . The motor  38  includes a rotor, generally indicated at  52 , configured toward an outside of the motor  38  and directly coupled to an input (not shown) of the drive member  36  and a stator, generally indicated at  42 , configured toward an inside of the motor  38  and positioned inside the rotor  52 . The drive member  36 , rotor  52 , and stator  42  are arranged substantially concentrically with each other. 
     More specifically, the rotor  52  and stator  42  combine with each other to make up the motor  38 . The actuator  30  defines a longitudinal axis and includes also the ground arm  46  that is configured to be connected to the wing rearward spar  28 . The actuator  30  includes also the output arm  48  that extends from the drive member  36 . In flight-control applications, the output arm  48  can define a hole  50  configured to receive a pin (not shown) that, in turn, is configured to be connected to an output-control surface (i.e., the aileron spar  18 ) of the aircraft. 
     As shown in  FIGS. 3 and 4 , in a version of the exemplary embodiment, the drive member  36  takes the form of a harmonic drive that includes a wave generator  40 . In particular, the harmonic drive is a gear of a gear train or set  36  having harmonic drive. However, it should be appreciated that the gearing can be other than harmonic. For example, the gear set  36  can be conventional (compound, planetary, simple, etc.). In any event, the gear set  36  acts as a speed-reduction device. 
     A reduction in number of components and, thereby, cost is achieved with design of the actuator  30 . More specifically, placement of the motor  38  within the gear or gear set  36  removes the drive shaft and one set of bearings of the known actuator and reduces inertia and number of parts of the actuator  30 . Also, the coupling and precision-machined housing of the known actuator are not required in the actuator  30  since an axis of rotation of the motor  38  is controlled by the gear set  36  itself. 
     “Reliability” analysis uses essentially a “reliability” factor for each component of a system multiplied by a number of components thereof. Thus, with fewer components of the same reliability with respect to each other, the system is more reliable. The actuator  30  has the fewest components for design of a motor/gear-set combination, leading to higher reliability of the actuator  30 . 
     The motor  38  is electric and can take the form of a brushless motor having the rotor  52  and stator  42 . The motor  38  is also frameless and of a high-performance type (i.e., has a high power-to-weight or power-to-volume ratio or power density). It should be appreciated that the motor  38  can be any suitable type of motor  38  that has a rotor  52  positioned on the outside. 
     The stator  42  is fixed and includes a plurality of coils  54 . An exterior/outer surface  52  of the rotor  52  acts as the wave generator  40  of the harmonic drive  36 . Alternatively, the wave generator  52  can be shaped to the exterior/outer surface. As shown in  FIG. 3 , an air gap  56  is defined between the rotor  52  and stator  42 . 
     The actuator  30  is configured to be employed with a thin-wing aircraft. Toward that end, arrangement of the actuator  30  does not have inefficiencies associated with packaging and is optimized for typical requirements of such an actuator in flight-control applications—power density, performance, and reliability. More specifically, the concentric packaging of the harmonic drive  36  and motor  38  (stator  42  and rotor  52 ) of the actuator  30  provides a higher power density. Also, a load path of the actuator  30  is a direct drive such that a drive shaft is not required, resulting in a lower inertia and, in turn, higher performance. Furthermore, the actuator  30  has few components (including removal of one set of bearings and no requirement as well for the compliant coupling or precision-machined housing), which lends itself to higher reliability and reduced cost. In addition, a total stack length of the actuator  30  can be changed to accommodate a higher output load, making the actuator  30  versatile for different applications. Moreover, the actuator  30  can achieve higher forces while it maintains a same cross-section thereof, making the actuator  30  versatile for the different applications. 
     While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions, or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various non-limiting embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.