Patent Publication Number: US-11650604-B2

Title: Yaw control systems for tailsitting biplane aircraft

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     The present application is a continuation of application Ser. No. 16/879,081 filed May 20, 2020. 
    
    
     TECHNICAL FIELD OF THE DISCLOSURE 
     The present disclosure relates, in general, to aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation and, in particular, to a tailsitting biplane aircraft having a counter-rotating coaxial rotor system with thrust vectoring capabilities. 
     BACKGROUND 
     Fixed-wing aircraft, such as airplanes, are capable of flight using wings that generate lift responsive to the forward airspeed of the aircraft, which is generated by thrust from one or more jet engines or propellers. The wings generally have an airfoil cross section that deflects air downward as the aircraft moves forward, generating the lift force to support the airplane in flight. Fixed-wing aircraft, however, typically require a runway that is hundreds or thousands of feet long for takeoff and landing. Unlike fixed-wing aircraft, vertical takeoff and landing (VTOL) aircraft do not require runways. Instead, VTOL aircraft are capable of taking off, hovering and landing vertically. One example of VTOL aircraft is a helicopter which is a rotorcraft having one or more rotors that provide lift to the aircraft. The rotors not only enable hovering and vertical takeoff and landing, but also enable forward, backward and lateral flight. These attributes make helicopters highly versatile for use in congested, isolated or remote areas where fixed-wing aircraft may be unable to takeoff and land. Helicopters, however, typically lack the forward airspeed and efficiency of fixed-wing aircraft. 
     A tiltrotor aircraft is another example of a VTOL aircraft. Tiltrotor aircraft generate lift and propulsion using proprotors that are typically coupled to nacelles mounted near the ends of a fixed wing. The nacelles rotate relative to the fixed wing such that the proprotors have a generally horizontal plane of rotation for vertical takeoff, hovering and landing and a generally vertical plane of rotation for forward flight, wherein the fixed wing provides lift and the proprotors provide forward thrust. In this manner, tiltrotor aircraft combine the vertical lift capability of a helicopter with the speed and range of fixed-wing aircraft. Tiltrotor aircraft, however, typically suffer from downwash inefficiencies during vertical takeoff and landing due to interference caused by the fixed wing. A further example of a VTOL aircraft is a tiltwing aircraft that features a rotatable wing that is generally horizontal for forward flight and rotates to a generally vertical orientation for vertical takeoff and landing. Propellers are coupled to the rotating wing to provide the required vertical thrust for takeoff and landing and the required forward thrust to generate lift from the wing during forward flight. The tiltwing design enables the slipstream from the propellers to strike the wing on its smallest dimension, thus improving vertical thrust efficiency as compared to tiltrotor aircraft. Tiltwing aircraft, however, are more difficult to control during hover as the vertically tilted wing provides a large surface area for crosswinds typically requiring tiltwing aircraft to have either cyclic rotor control or an additional thrust station to generate a moment. 
     SUMMARY 
     In one aspect, the present disclosure is directed to an aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation. The aircraft includes an airframe with a first wing, a second wing and a fuselage that extends between the first and second wings. A propulsion assembly is coupled to the fuselage and includes a counter-rotating coaxial rotor system that is tiltable relative to the fuselage to generate a thrust vector. First and second yaw vanes extend aftwardly from the fuselage. A flight control system is configured to direct the thrust vector of the coaxial rotor system and control movements of the yaw vanes. In the VTOL orientation, differential operation of the yaw vanes provides yaw authority for the aircraft. In the biplane orientation, collective operation of the yaw vanes provides yaw authority for the aircraft. 
     In certain embodiments, the propulsion assembly may include a motor assembly such that the coaxial rotor system and the motor assembly are tiltable relative to the fuselage to generate the thrust vector. In some embodiments, the coaxial rotor system may be configured to provide thrust in line with a yaw axis of the aircraft, in the VTOL orientation, and in line with a roll axis of the aircraft, in the biplane orientation. In certain embodiments, in the VTOL orientation, the first wing is forward of the fuselage and the second wing is aft of the fuselage and, in the biplane orientation, the first wing is below the fuselage and the second wing is above the fuselage. In some embodiments, the first and second yaw vanes may be pivotably coupled to the fuselage. 
     In certain embodiments, in the VTOL orientation, the aircraft has a VTOL yaw axis that extends through the fuselage and the propulsion assembly and, in the biplane orientation, the aircraft has a biplane yaw axis that extends through the fuselage and the wings. In such embodiments, in the VTOL orientation, the first and second yaw vanes may be symmetrically disposed relative to the VTOL yaw axis and, in the biplane orientation, the first and second yaw vanes may be aftwardly disposed relative to the biplane yaw axis. In some embodiments, in the VTOL orientation, the first yaw vane may be actuated in a first direction and the second yaw vane may be actuated in a second direction that is opposite of the first direction to provide yaw authority for the aircraft. In certain embodiments, in the VTOL orientation, downwash generated by the propulsion assembly acting on the actuated first and second yaw vanes may create first and second yaw moments that urge the aircraft to rotate about the VTOL yaw axis. In some embodiments, in the biplane orientation, the first and second yaw vanes may both be actuated in a first direction to provide yaw authority for the aircraft. In certain embodiments, in the biplane orientation, airflow generated responsive to the forward airspeed of the aircraft acting on the actuated first and second yaw vanes may create yaw moments that urges the aircraft to rotate about the biplane yaw axis. 
     In another aspect, the present disclosure is directed to an aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation. The aircraft includes an airframe with a first wing, a second wing and a fuselage that extends between the first and second wings. A propulsion assembly is coupled to the fuselage and includes a counter-rotating coaxial rotor system having first and second rotor assemblies. The coaxial rotor system is tiltable relative to the fuselage to generate a thrust vector. First and second yaw vanes extend aftwardly from the fuselage. A flight control system is configured to direct the thrust vector of the coaxial rotor system, control rotor speeds and collective pitches of the first and second rotor assemblies and control movements of the yaw vanes. In the VTOL orientation, differential operation of the first and second rotor assemblies and differential operation of the yaw vanes provide yaw authority for the aircraft. 
     In some embodiments, in the VTOL orientation, the flight control system may be configured to provide yaw authority for the aircraft responsive to actuating the first yaw vane in a first direction, actuating the second yaw vane in a second direction that is opposite of the first direction and differentially operating the first and second rotor assemblies by adjusting the rotor speeds. In certain embodiments, in the VTOL orientation, the flight control system may be configured to provide yaw authority for the aircraft responsive to actuating the first yaw vane in a first direction, actuating the second yaw vane in a second direction that is opposite of the first direction and differentially operating the first and second rotor assemblies by adjusting the collective pitches. In some embodiments, in the VTOL orientation, the flight control system may be configured to provide yaw authority for the aircraft responsive to actuating the first yaw vane in a first direction, actuating the second yaw vane in a second direction that is opposite of the first direction and differentially operating the first and second rotor assemblies by adjusting the rotor speeds and the collective pitches. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       For a more complete understanding of the features and advantages of the present disclosure, reference is now made to the detailed description along with the accompanying figures in which corresponding numerals in the different figures refer to corresponding parts and in which: 
         FIGS.  1 A- 1 G  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIGS.  2 A- 2 I  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in a sequential flight operating scenario in accordance with embodiments of the present disclosure; 
         FIGS.  3 A- 3 F  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in a sequential flight operating scenario in accordance with embodiments of the present disclosure; 
         FIGS.  4 A- 4 D  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in various flight configurations in accordance with embodiments of the present disclosure; 
         FIGS.  5 A- 5 H  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in various flight configurations in accordance with embodiments of the present disclosure; 
         FIG.  6    depicts a propulsion assembly and gimbal assembly for a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIG.  7    depicts an aft door of a fuselage, in partial cutaway, for a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIGS.  8 A- 8 C  depict a tail assembly in various operating configurations for a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIG.  9 A  is a systems diagram of one implementation of a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIG.  9 B  is a block diagram of autonomous and remote control systems for a tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure; 
         FIGS.  10 A- 10 D  are schematic illustrations of a tailsitting biplane aircraft having a coaxial rotor system in various roles in accordance with embodiments of the present disclosure; and 
         FIGS.  11 A- 11 D  are schematic illustrations of various tailsitting biplane aircraft having a coaxial rotor system in accordance with embodiments of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
     While the making and using of various embodiments of the present disclosure are discussed in detail below, it should be appreciated that the present disclosure provides many applicable inventive concepts, which can be embodied in a wide variety of specific contexts. The specific embodiments discussed herein are merely illustrative and do not delimit the scope of the present disclosure. In the interest of clarity, not all features of an actual implementation may be described in the present disclosure. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer&#39;s specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming but would be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure. 
     In the specification, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of the present disclosure, the devices, members, apparatuses, and the like described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction. As used herein, the term “coupled” may include direct or indirect coupling by any means, including moving and/or non-moving mechanical connections. 
     Referring to  FIGS.  1 A- 1 G  in the drawings, a tailsitting biplane aircraft operable to transition between thrust-borne lift in a VTOL orientation and wing-borne lift in a biplane orientation is schematically illustrated and generally designated as aircraft  10 .  FIGS.  1 A,  1 C and  1 E  depict aircraft  10  in the VTOL orientation wherein the propulsion assembly provides thrust-borne lift enabling aircraft  10  to accomplish vertical takeoffs, hover and vertical landings.  FIGS.  1 B,  1 D and  1 F  depict aircraft  10  in the biplane orientation wherein the propulsion assembly provides forward thrust with the forward airspeed of aircraft  10  providing wing-borne lift enabling aircraft  10  to have a high speed, high efficiency and/or high endurance forward flight mode. Aircraft  10  has a longitudinal axis  10   a  that may be referred to as the roll axis, a lateral axis  10   b  that may be referred to as the pitch axis and a vertical axis  10   c  that may be referred to as the yaw axis, as best seen in  FIGS.  1 E and  1 F . When longitudinal axis  10   a  and lateral axis  10   b  are both in a horizontal plane and normal to the local vertical in the earth&#39;s reference frame, aircraft  10  has a level flight attitude. 
     In the illustrated embodiment, aircraft  10  has an airframe  12  including wings  14 ,  16  each having an airfoil cross-section that generates lift responsive to the forward airspeed of aircraft  10 , in the biplane orientation. Wings  14 ,  16  may be formed as single members or may be formed from multiple wing sections. The outer skins of wings  14 ,  16  are preferably formed from high strength and lightweight materials such as fiberglass, carbon, plastic, metal or other suitable material or combination of materials. As best seen in  FIG.  1 B , wing  14  has a dihedral configuration and wing  16  has an anhedral configuration. In the illustrated embodiment, wing  14  has a dihedral angle of about twelve degrees and wing  16  has an anhedral angle of about twelve degrees. In other embodiments, wing  14  may have a dihedral angle of between eight degrees and sixteen degrees or another suitable dihedral angle. Likewise, wing  16  may have an anhedral angle between eight degrees and sixteen degrees or other suitable anhedral angle. The dihedral and anhedral configuration of wings  14 ,  16  provide enhanced ground stability of aircraft  10 , while the dual wing design provides a compact footprint on the ground. 
     As best seen in  FIG.  1 E , wings  14 ,  16  have swept wing designs. In the illustrated embodiment, wings  14 ,  16  have a quarter chord sweep angle between fifteen degrees and thirty degrees such as a quarter chord sweep angle between twenty degrees and twenty-five degrees or a quarter chord sweep angle of about twenty-two degrees. In the illustrated embodiment, the leading edge sweep angle is greater than the quarter chord sweep angle and is about twenty-five degrees, the half chord sweep angle is less than the quarter chord sweep angle and is about nineteen degrees and the trailing edge sweep angle is less than the half chord sweep angle and is about twelve and half degrees. As illustrated, the sweep angle progressively decreases from the leading edge to the trailing edge forming a tapered swept wing design. In other embodiments, the sweep angle may remain constant from the leading edge to the trailing edge forming a simple swept wing design, the leading edge may have a sweep angle and the trailing edge may not have a sweep angle forming a delta swept wing design or the leading edge may have a positive sweep angle and the trailing edge may have a negative sweep angle forming a trapezoidal swept wing design. 
     Airframe  12  also includes a fuselage  18  that extends generally perpendicularly between wings  14 ,  16 . Fuselage  18  is preferably formed from high strength and lightweight materials such as fiberglass, carbon, plastic, metal or other suitable material or combination of materials. As best seen in  FIG.  1 G , fuselage  18  has an aerodynamic shape to minimize drag during high speed forward flight. In addition, fuselage  18  preferably has a length in the longitudinal direction configured to minimize interference drag between wings  14 ,  16 . For example, the longitudinal length of fuselage  18  may have a ratio to the wingspan of wings  14 ,  16  of between 1 to 2 and 1 to 3 such as a ratio of about 1 to 2.5. Fuselage  18  has an upper flange  18   a  configured to receive wing  14  and a lower flange  18   b  configured to receive wing  16 . In the illustrated embodiment, wing  14  is aft of fuselage  18  and wing  16  is forward of fuselage  18  in the VTOL orientation and wing  14  is above fuselage  18  and wing  16  is below fuselage  18  in the biplane orientation. Wings  14 ,  16  may be attachable to and detachable from fuselage  18  and may be standardized and/or interchangeable units and preferably line replaceable units providing easy installation and removal from fuselage  18 . The use of line replaceable wings is beneficial in maintenance situations if a fault is discovered with a wing. In this case, the faulty wing can be decoupled from fuselage  18  by simple operations and another wing can then be attached to fuselage  18 . In other embodiments, wings  14 ,  16  may be permanently coupled to fuselage  18 . In either case, the connections between wings  14 ,  16  and fuselage  18  are preferably stiff connections. In the illustrated embodiment, fuselage  18  includes an aft door  18   c  that is pivotably coupled to main body  18   d  of fuselage  18 . Aft door  18   c  may be rotated relative to main body  18   d  to allow access to the inside of fuselage  18 , for cargo, passengers, crew or the like. During flight operations, aft door  18   c  is secured to main body  18   d  to prevent relative rotation therebetween. Operation of aft door  18   c  between the open and closed positions may be manual or automated. In other embodiments, an aft door for fuselage  18  may be formed from a pair of clamshell doors each of which is rotatable relative to main body  18   d  to allow access to the inside of fuselage  18 . 
     As best seen in  FIGS.  1 C and  1 D , fuselage  18  houses a flight control system  20  of aircraft  10 . Flight control system  20  is preferably a redundant digital flight control system including multiple independent flight control computers. For example, the use of a triply redundant flight control system  20  improves the overall safety and reliability of aircraft  10  in the event of a failure in flight control system  20 . Flight control system  20  preferably includes non-transitory computer readable storage media including a set of computer instructions executable by one or more processors for controlling the operation of aircraft  10 . Flight control system  20  may be implemented on one or more general-purpose computers, one or more special purpose computers or other machines with memory and processing capability. For example, flight control system  20  may include one or more memory storage modules including, but not limited to, internal storage memory such as random access memory, non-volatile memory such as read only memory, removable memory such as magnetic storage memory, optical storage, solid-state storage memory or other suitable memory storage entity. Flight control system  20  may be a microprocessor-based system operable to execute program code in the form of machine-executable instructions. In addition, flight control system  20  may be selectively connectable to other computer systems via a proprietary encrypted network, a public encrypted network, the Internet or other suitable communication network that may include both wired and wireless connections. 
     Aircraft  10  includes a propulsion assembly  22  that is coupled to fuselage  18 . Propulsion assemblies such as propulsion assembly  22  may be attachable to and detachable from fuselage  18  and may be standardized and/or interchangeable units and preferably line replaceable units providing easy installation and removal from fuselage  18 . The use of line replaceable propulsion assemblies is beneficial in maintenance situations if a fault is discovered with a propulsion assembly. In this case, the faulty propulsion assembly can be decoupled from fuselage  18  by simple operations and another propulsion assembly can then be attached to fuselage  18 . In the illustrated embodiment, propulsion assembly  22  includes a counter-rotating coaxial rotor system  24  that has first and second rotors assemblies  24   a ,  24   b  that share a common axis of rotation and counter-rotate relative to one another. First and second rotors assemblies  24   a ,  24   b  may be referred to as upper rotor assembly  24   a  and lower rotor assembly  24   b  in the VTOL orientation of aircraft  10  and as forward rotor assembly  24   a  and aft rotor assembly  24   b  in the biplane orientation of aircraft  10 . 
     In the VTOL orientation of aircraft  10 , when coaxial rotor system  24  is not thrust vectoring, upper rotor assembly  24   a  and lower rotor assembly  24   b  rotate about yaw or vertical axis  10   c , as best seen in  FIG.  1 E , providing thrust in line with the center of gravity of aircraft  10  and in line with yaw axis  10   c . In the biplane orientation of aircraft  10 , when coaxial rotor system  24  is not thrust vectoring, forward rotor assembly  24   a  and aft rotor assembly  24   b  rotate about roll or longitudinal axis  10   a , as best seen in  FIG.  1 F , providing thrust in line with the center of gravity of aircraft  10  and in line with roll axis  10   a . In the illustrated embodiment, rotors assemblies  24   a ,  24   b  are rigid rotor systems and/or hingeless rotor systems. In other embodiments, rotors assemblies  24   a ,  24   b  could have alternate rotor system designs such as fully or partially articulated rotor systems. As illustrated in  FIG.  1 E , in the VTOL orientation, yaw axis  10   c  extends through fuselage  18  and propulsion assembly  22  and may be referred to as VTOL yaw axis  10   c . Likewise, roll axis  10   a  extends through fuselage  18  and wings  14 ,  16  and may be referred to as VTOL roll axis  10   a . As illustrated in  FIG.  1 F , in the biplane orientation, yaw axis  10   c  extends through fuselage  18  and wings  14 ,  16  and may be referred to as biplane yaw axis  10   c . Likewise, roll axis  10   a  extends through fuselage  18  and propulsion assembly  22  and may be referred to as biplane roll axis  10   a.    
     As best seen in  FIG.  1 B , forward rotor assembly  24   a  has four rotor blades and aft rotor assembly  24   b  has four rotor blades with forward rotor assembly  24   a  rotating counterclockwise, as indicated by arrow  24   c , and with aft rotor assembly  24   b  rotating clockwise, as indicated by arrow  24   d , when aircraft  10  is viewed from the front. In the illustrated embodiment, each rotor blade has a root to tip twist between forty degrees and fifty degrees such as a root to tip twist between forty-five degrees and forty-eight degrees or a root to tip twist of about forty-seven degrees. Forward rotor assembly  24   a  and aft rotor assembly  24   b  may have the same or similar diameters or may have different diameters such as forward rotor assembly  24   a  having a larger diameter than aft rotor assembly  24   b . In the illustrated embodiment, the rotor disk of coaxial rotor system  24  has a ratio to the wingspan of wings  14 ,  16  of between 1 to 1 and 1 to 3 such as a ratio of about 1 to 2.5. In other embodiments, the rotor disk of coaxial rotor system  24  could have a ratio to the wingspan of wings  14 ,  16  of greater than 1 to 1 or less than 1 to 3. 
     In the illustrated embodiment, aircraft  10  is an electric vertical takeoff and landing (eVTOL) aircraft having two electric motors depicted as a motor assembly  22   a , with each of the electric motors driving one of forward rotor assembly  24   a  and aft rotor assembly  24   b . In other embodiments, aircraft  10  may have a single electric motor driving both forward rotor assemblies  24   a ,  24   b  or may have one or more internal combustion engines driving rotor assemblies  24   a ,  24   b  via one or more suitable transmissions. In the illustrated embodiment, forward rotor assembly  24   a  and aft rotor assembly  24   b  are independently controllable rotor assemblies configured for independent variable speed control and independent collective pitch control with no cyclic pitch control. In other embodiments, rotor assemblies  24   a ,  24   b  could operate at a constant speed, could have fixed pitch rotor blades and/or could have cyclic pitch control. In the illustrated embodiment, propulsion assembly  22  including coaxial rotor system  24  and motor assembly  22   a  is tiltable relative to fuselage  18  to provide omnidirectional thrust vectoring capability to aircraft  10  with the counter rotation of rotor assemblies  24   a ,  24   b  cancelling the gyroscopic moments. For example, propulsion assembly  22  may be gimbal mounted to fuselage  18  via propulsion assembly housing  22   b , which is part of airframe  12 , such that coaxial rotor system  24  and motor assembly  22   a  tilt about orthogonal pivot axes such as a pitch pivot axis and a lateral pivot axis. In other embodiments, propulsion assembly  22  may be gimbal mounted to fuselage  18  such that coaxial rotor system  24  tilts about two non-orthogonal pivot axes or such that coaxial rotor system  24  tilts about only one pivot axis such as the pitch pivot axis. In still other embodiments, coaxial rotor system  24  may be tiltable relative to fuselage  18  with motor assembly  22   a  being non-tiltable. 
     Aircraft  10  has a distributed array of control surfaces carried by tail assemblies  26   a ,  26   b ,  26   c ,  26   d , which may collectively be referred to as tail assemblies  26 . In the illustrated embodiment, tail assemblies  26   a ,  26   b  are respectively coupled to wingtips  14   a ,  14   b  of wing  14  and tail assemblies  26   c ,  26   d  are respectively coupled to wingtips  16   a ,  16   b  of wing  16  such that tail assemblies  26   a ,  26   b ,  26   c ,  26   d  are positioned outboard of the rotor disk of coaxial rotor system  24 . Tail assemblies  26  may be independently attachable to and detachable from the wingtips and may be standardized and/or interchangeable units and preferably line replaceable units providing easy installation and removal from the wingtips. The use of line replaceable tail assemblies is beneficial in maintenance situations if a fault is discovered with one of the tail assemblies. In this case, the faulty tail assembly can be decoupled from the wingtip by simple operations and another tail assembly can then be attached to the wingtip. In other embodiments, tail assemblies  26  may be permanently coupled to wings  14 ,  16 . 
     Tail assembly  26   a  includes a pair of aerosurfaces depicted as a vertical stabilizer  28   a  and an elevon  30   a . Tail assembly  26   b  includes a pair of aerosurfaces depicted as a vertical stabilizer  28   b  and an elevon  30   b . Tail assembly  26   c  includes a pair of aerosurfaces depicted as a vertical stabilizer  28   c  and an elevon  30   c . Tail assembly  26   d  includes a pair of aerosurfaces depicted as a vertical stabilizer  28   d  and an elevon  30   d . Vertical stabilizers  28   a ,  28   b ,  28   c ,  28   d  may collectively be referred to as vertical stabilizers  28  and elevons  30   a ,  30   b ,  30   c ,  30   d  may collectively be referred to as elevons  30 . In the illustrated embodiment, vertical stabilizers  28  are fixed aerosurfaces. In other embodiments, vertical stabilizers  28  could operate as rudders. In the illustrated embodiment, elevons  30  are pivoting aerosurfaces that are rotatable about respective elevon axes that may be generally parallel with wings  14 ,  16  at the respective dihedral and anhedral angles. When operated collectively, elevons  30  serve as elevators to control the pitch or angle of attack of aircraft  10 , in the biplane orientation. When operated differentially, elevons  30  serve as ailerons to control the roll or bank of aircraft  10 , in the biplane orientation. 
     Aircraft  10  includes a plurality of electrical power sources depicted as batteries  32   a ,  32   b ,  32   c ,  32   d , which may collectively be referred to as batteries  32 . In the illustrated embodiment, batteries  32  form a distributed power system in which each battery  32   a ,  32   b ,  32   c ,  32   d  is located in a receiving pocket of one of the tail assemblies  26   a ,  26   b ,  26   c ,  26   d  such that batteries  32  provide inertial relief to wings  14 ,  16 . Batteries  32  provide power to flight control system  20 , propulsion assembly  22  and other power consumers via a power management system including, for example, a centralized DC bus. Alternatively or additionally, batteries may be housed within fuselage  18  and/or wings  14 ,  16 . In some embodiments, aircraft  10  may have a hybrid power system that includes one or more internal combustion engines and an electric generator. Preferably, the electric generator is used to charge batteries  32 . In other embodiments, the electric generator may provide power directly to the power management system and/or the power consumers such as propulsion assembly  22 . In still other embodiments, aircraft  10  may use fuel cells as the electrical power sources. The fuel cell may be located in the receiving pockets of tail assemblies  26 , in fuselage  18  and/or in wings  14 ,  16 . 
     Aircraft  10  includes a pair of yaw vanes  34   a ,  34   b  that are pivotably coupled to an aft end of fuselage  18 . Yaw vanes  34   a ,  34   b  may be operated differentially to generate yaw moments when aircraft  10  is in the VTOL orientation and may be operated collectively to generate yaw moments when aircraft  10  is in the biplane orientation. Aircraft  10  has a plurality of landing gear assemblies  36   a ,  36   b ,  36   c ,  36   d  that may collectively be referred to as landing gear assemblies  36 . Landing gear assemblies  36   a ,  36   b ,  36   c ,  36   d  are positioned at the distal end of respective tail assemblies  26   a ,  26   b ,  26   c ,  26   d . The landing gear assemblies  36  may be passively operated pneumatic landing struts or actively operated telescoping landing struts. In other embodiments, landing gear assemblies  36  may include wheels that enable aircraft  10  to taxi and perform other ground maneuvers. In such embodiments, landing gear assemblies  36  may provide a passive brake system or may include active brakes such as an electromechanical braking system or a manual braking system to facilitate parking during ground operations. 
     Aircraft  10  may be a manned or unmanned aircraft. Flight control system  20  may autonomously control some or all aspects of flight operations for aircraft  10 . Flight control system  20  is also operable to communicate with remote systems, such as a ground station via a wireless communications protocol. The remote system may be operable to receive flight data from and provide commands to flight control system  20  to enable remote flight control over some or all aspects of flight operations for aircraft  10 . The remote flight control and/or autonomous flight control may be augmented or supplanted by onboard pilot flight control during manned missions. Regardless of the input, aircraft  10  preferably utilizes a fly-by-wire system that transmits electronic signals from flight control system  20  to the actuators of controlled systems to control the flight dynamics of aircraft  10  including controlling the movements of elevons  30 , yaw vanes  34   a ,  34   b  and propulsion assembly  22 . Flight control system  20  communicates with the controlled systems via a fly-by-wire communications network within airframe  12 . In addition, flight control system  20  receives data from a plurality of sensors  40  such as one or more position sensors, attitude sensors, speed sensors, altitude sensors, heading sensors, environmental sensors, fuel sensors, temperature sensors, location sensors and the like to enhance flight control capabilities. 
     Referring additionally to  FIGS.  2 A- 2 I  in the drawings, a sequential flight-operating scenario of aircraft  10  will now be described. As best seen in  FIG.  2 A , aircraft  10  is in a tailsitter position on a surface such as the ground or the deck of an aircraft carrier. When aircraft  10  is ready for a mission, flight control system  20  commences operations providing flight commands to the various systems of aircraft  10 . Flight control system  20  may be operating responsive to autonomous flight control, remote flight control, onboard pilot flight control or a combination thereof. For example, it may be desirable to utilize onboard pilot or remote flight control during certain maneuvers such as takeoffs and landings but rely on autonomous flight control during hover, high speed forward flight and transitions between wing-borne flight and thrust-borne flight. 
     As best seen in  FIG.  2 B , aircraft  10  has performed a vertical takeoff and is engaged in thrust-borne lift in the VTOL orientation of aircraft  10 . As illustrated, upper rotor assembly  24   a  and lower rotor assembly  24   b  are counter-rotating in generally parallel horizontal planes. As longitudinal axis  10   a  and lateral axis  10   b  (denoted as the target) are both in a horizontal plane H that is normal to the local vertical in the earth&#39;s reference frame, aircraft  10  has a level flight attitude. In the VTOL orientation, wing  16  is the forward wing and wing  14  is the aft wing. Flight control system  20  independently controls and operates upper rotor assembly  24   a  and lower rotor assembly  24   b  including independently controlling rotor speed and collective pitch. In addition, flight control system  20  controls the tilt of propulsion assembly  22  relative to fuselage  18  to generate a thrust vector. 
     During hover, flight control system  20  may utilize speed control and/or collective pitch control of upper rotor assembly  24   a  and lower rotor assembly  24   b  to cause aircraft  10  to climb, descend or maintain a stable hover. Also during hover, flight control system  20  may utilize thrust vectoring of propulsion assembly  22  to provide translation authority for aircraft  10 . For example, as best seen in  FIG.  4 A , propulsion assembly  22  is tiltable forward and aftward relative to fuselage  18  to provide translation authority to aircraft  10  in the fore/aft direction, as indicated by arrow  42 . When propulsion assembly  22  is tilted aftward relative to fuselage  18 , as indicated by dotted propulsion assembly  22   c , propulsion assembly  22  generates a thrust vector having a vertical component  44  providing thrust-borne lift for aircraft  10  and an aftward component  46  that urges aircraft  10  to translate in the aftward direction. When propulsion assembly  22  is tilted forward relative to fuselage  18 , as indicated by dotted propulsion assembly  22   d , propulsion assembly  22  generates a thrust vector having vertical component  44  providing thrust-borne lift for aircraft  10  and a forward component  48  that urges aircraft  10  to translate in the forward direction. 
     As another example, as best seen in  FIG.  4 B , propulsion assembly  22  is tiltable to the right and to the left relative to fuselage  18  to provide translation authority to aircraft  10  in the lateral direction, as indicated by arrow  50 . When propulsion assembly  22  is tilted to the right relative to fuselage  18 , as indicated by dotted propulsion assembly  22   e , propulsion assembly  22  generates a thrust vector having a vertical component  52  providing thrust-borne lift for aircraft  10  and a lateral component  54  that urges aircraft  10  to translate to the right. When propulsion assembly  22  is tilted left relative to fuselage  18 , as indicated by dotted propulsion assembly  22   f , propulsion assembly  22  generates a thrust vector having vertical component  52  providing thrust-borne lift for aircraft  10  and a lateral component  56  that urges aircraft  10  to translate to the left. 
     In the illustrated embodiment, the thrust vectoring capability of propulsion assembly  22  is achieved by operating a gimbal assembly  60 , as best seen in  FIG.  6   . Gimbal assembly  60  includes an inner gimbal ring  62  that is coupled to motor assembly  22   a  that includes electric motors  22   g ,  22   h  that respectively provide torque and rotational energy to upper rotor assembly  24   a  and lower rotor assembly  24   b . Gimbal assembly  60  also includes an outer gimbal ring  66  that is rotatably coupled to propulsion assembly housing  22   b  and rotatably coupled to inner gimbal ring  62 . An inner gimbal ring actuator  70  is configured to tilt inner gimbal ring  62  relative to outer gimbal ring  66  about pitch pivot axis  72  via linkage  74 , responsive to commands from flight control system  20 . An outer gimbal ring actuator  76  is configured to tilt outer gimbal ring  66  relative to propulsion assembly housing  22   b  about lateral pivot axis  78  via linkage  80 , responsive to commands from flight control system  20 . In this manner, propulsion assembly  22  including coaxial rotor system  24  and motor assembly  22   a  is tilted relative to fuselage  18  to generate the thrust vector. Even though aircraft  10  has been depicted in  FIG.  4 A  and described in reference thereto as being configurable for fore/aft translation and even though aircraft  10  has been depicted in  FIG.  4 B  and described in reference thereto as being configurable for lateral translation, it should be understood by those having ordinary skill in the art that the orthogonal pivot axes of gimbal assembly  60  provide for tilting of propulsion assembly  22  in any radial direction relative to fuselage  18  such that propulsion assembly  22  has omnidirectional thrust vectoring capability and such that aircraft  10  has omnidirectional translation capability in hover. 
     Continuing with the sequential flight-operating scenario, aircraft  10  remains in the hover operation in  FIG.  2 B . During hover, flight control system  20  may utilize differential rotor speed, differential collective pitch and/or differential yaw vane positioning to provide yaw authority for aircraft  10 . For example, to maintain a stable hover, differential rotor speed may be used wherein the average rotor speed of upper rotor assembly  24   a  and lower rotor assembly  24   b  is held constant while increasing the rotor speed of one of upper rotor assembly  24   a  and lower rotor assembly  24   b  and decreasing the rotor speed of the other of upper rotor assembly  24   a  and lower rotor assembly  24   b  to create a torque imbalance that provides yaw authority for aircraft  10 . In the illustrated embodiment with upper rotor assembly  24   a  rotating counterclockwise and lower rotor assembly  24   b  rotating clockwise when aircraft  10  is viewed from above, increasing the rotor speed of upper rotor assembly  24   a  and decreasing the rotor speed of lower rotor assembly  24   b  will cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction. Similarly, decreasing the rotor speed of upper rotor assembly  24   a  and increasing the rotor speed of lower rotor assembly  24   b  will cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction. 
     As another example, to maintain a stable hover, differential collective pitch may be used wherein the effective collective pitch of upper rotor assembly  24   a  and lower rotor assembly  24   b  is held constant while increasing the collective pitch of one of upper rotor assembly  24   a  and lower rotor assembly  24   b  and decreasing the collective pitch of the other of upper rotor assembly  24   a  and lower rotor assembly  24   b  to create a torque imbalance that provides yaw authority for aircraft  10 . In the illustrated embodiment with upper rotor assembly  24   a  rotating counterclockwise and lower rotor assembly  24   b  rotating clockwise when aircraft  10  is viewed from above, increasing the collective pitch of upper rotor assembly  24   a  and decreasing the collective pitch of lower rotor assembly  24   b  will cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction. Similarly, decreasing the collective pitch of upper rotor assembly  24   a  and increasing the collective pitch of lower rotor assembly  24   b  will cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction. 
     In a further example, yaw vanes  34   a ,  34   b  may be operated differentially to create yaw moments in response to propulsion downwash generated by propulsion system  22  over yaw vanes  34   a ,  34   b . As best seen in  FIG.  4 C , when yaw vane  34   a  is shifted to the left and yaw vane  34   b  is shifted to the right, when aircraft  10  is viewed from a forward position during hover, the propulsion downwash acting on yaw vanes  34   a ,  34   b  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, as indicated by arrow  90 . Similarly, as best seen in  FIG.  4 D , when yaw vane  34   a  is shifted to the right and yaw vane  34   b  is shifted to the left, when aircraft  10  is viewed from a forward position during hover, the propulsion downwash acting on yaw vanes  34   a ,  34   b  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, as indicated by arrow  92 . In the illustrated embodiment, the differential positioning of yaw vanes  34   a ,  34   b  is achieved by operation of actuators positioned within aft door  18   c  of fuselage  18 . 
     As best seen in  FIG.  7   , yaw vanes  34   a ,  34   b  are symmetrically disposed and pivotably coupled to the aft end of fuselage  18  and more specifically to the aft end of aft door  18   c . It is noted that in the illustrated embodiment, a portion of aft door  18   c  has been cut away to reveal yaw vane actuators  94   a ,  94   b  that are mounted within aft door  18   c . Yaw vane actuator  94   a  is configured to tilt yaw vane  34   a  relative to aft door  18   c  about a pivot axis  96  via linkage  98   a  that extends through a slot  18   d  in aft door  18   c . Yaw vane actuator  94   b  is configured to tilt yaw vane  34   b  relative to aft door  18   c  about pivot axis  96  via linkage  98   b  that extends through a slot  18   f  in aft door  18   c . Yaw vane actuators  94   a ,  94   b  are independently operated responsive to commands from flight control system  20  such that yaw vanes  34   a ,  34   b  may be collectively or differentially pivoted relative to aft door  18   c.    
     In addition to using the yaw authority mechanisms described herein to individually provide yaw authority for aircraft  10  during hover, flight control system  20  can command multiple yaw authority mechanisms to operate together to provide yaw authority for aircraft  10  during hover. For example, to cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, flight control system  20  could increase the rotor speed and increase the collective pitch of upper rotor assembly  24   a  while decreasing the rotor speed and decreasing the collective pitch of lower rotor assembly  24   b . To cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, flight control system  20  could decrease the rotor speed and decrease the collective pitch of upper rotor assembly  24   a  while increasing the rotor speed and increasing the collective pitch of lower rotor assembly  24   b . As another example, to cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, flight control system  20  could increase the rotor speed of upper rotor assembly  24   a , decrease the rotor speed of lower rotor assembly  24   b , shift yaw vane  34   a  to the left and shift yaw vane  34   b  to the right. To cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, flight control system  20  could decrease the rotor speed of upper rotor assembly  24   a , increase the rotor speed of lower rotor assembly  24   b , shift yaw vane  34   a  to the right and shift yaw vane  34   b  to the left. 
     In a further example, to cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, flight control system  20  could increase the collective pitch of upper rotor assembly  24   a , decrease the collective pitch of lower rotor assembly  24   b , shift yaw vane  34   a  to the left and shift yaw vane  34   b  to the right. To cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, flight control system  20  could decrease the collective pitch of upper rotor assembly  24   a , increase the collective pitch of lower rotor assembly  24   b , shift yaw vane  34   a  to the right and shift yaw vane  34   b  to the left. Additionally, to cause aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, flight control system  20  could increase the collective pitch and rotor speed of upper rotor assembly  24   a , decrease the collective pitch and rotor speed of lower rotor assembly  24   b , shift yaw vane  34   a  to the left and shift yaw vane  34   b  to the right. To cause aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, flight control system  20  could decrease the collective pitch and rotor speed of upper rotor assembly  24   a , increase the collective pitch and rotor speed of lower rotor assembly  24   b , shift yaw vane  34   a  to the right and shift yaw vane  34   b  to the left. Using more than one and/or different combinations of yaw authority mechanisms can be beneficial depending upon aircraft parameters, flight dynamics and/or environmental factors including altitude, attitude, temperature, thrust to weight ratio, wind speed, wind direction, desired yaw rate and other considerations known to those having ordinary skill in the art. 
     In embodiments wherein the rotor disk of coaxial rotor system  24  has a ratio to the wingspan of wings  14 ,  16  on the order of 1 to 1 or greater, differential operations of elevons  30  may be used to complement other yaw authority mechanisms in hover or as a standalone yaw authority mechanism in hover. For example, when elevons  30   a ,  30   c  are tilted forward (see  FIG.  8 B ) and elevons  30   b ,  30   d  are tilted aftward (see  FIG.  8 C ), propulsion downwash generated by propulsion system  22  over elevons  30  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in the counterclockwise direction, as seen from above in  FIG.  1 E . Similarly, when elevons  30   b ,  30   d  are tilted forward (see  FIG.  8 B ) and elevons  30   a ,  30   c  are tilted aftward (see  FIG.  8 C ), propulsion downwash generated by propulsion system  22  over elevons  30  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in the clockwise direction, as seen from above in  FIG.  1 E . 
     In embodiments wherein propulsion assembly  22  is gimbal mounted to fuselage  18  with a single axis gimbal in which propulsion assembly  22  is tiltable only forward and aftward relative to fuselage  18 , pitch axis thrust vectoring provides translation authority to aircraft  10  in the fore/aft direction  42  in hover (see  FIG.  4 A ). When propulsion assembly  22  is tilted aftward relative to fuselage  18 , as indicated by dotted propulsion assembly  22   c , propulsion assembly  22  generates a thrust vector having a vertical component  44  providing thrust-borne lift for aircraft  10  and an aftward component  46  that urges aircraft  10  to translate in the aftward direction. When propulsion assembly  22  is tilted forward relative to fuselage  18 , as indicated by dotted propulsion assembly  22   d , propulsion assembly  22  generates a thrust vector having vertical component  44  providing thrust-borne lift for aircraft  10  and a forward component  48  that urges aircraft  10  to translate in the forward direction. In such single axis gimbal embodiments of propulsion assembly  22 , translation authority for aircraft  10  in the lateral direction is provided by collective operation of yaw vanes  34   a ,  34   b  to create lateral forces acting on yaw vanes  34   a ,  34   b  in response to propulsion downwash generated by propulsion system  22  over yaw vanes  34   a ,  34   b . Collectively shifting yaw vanes  34   a ,  34   b  to the left, urges aircraft  10  to translate to the right (see  FIG.  1 B ). Likewise, collectively shifting yaw vanes  34   a ,  34   b  to the right, urges aircraft  10  to translate to the left. Coordinated pitch axis thrust vectoring and collective yaw vane operation provide omnidirectional translation capability to aircraft  10  in hover. Alternatively or additionally, in such single axis gimbal embodiments of propulsion assembly  22 , aircraft  10  is operable to translate in any direction by first, rotating aircraft  10  about yaw axis  10   c  to a desired fore/aft or longitudinal orientation then second, pitch axis thrust vectoring to translate in the desired direction. 
     Continuing with the sequential flight-operating scenario, aircraft  10  has completed the vertical ascent to a desired elevation in  FIG.  2 C  and may now begin the transition from thrust-borne lift to wing-borne lift. As best seen from the progression of  FIGS.  2 C- 2 E , aircraft  10  is operable to pitch down from the VTOL orientation toward the biplane orientation to enable high speed and/or long range forward flight. As seen in  FIG.  2 C , aircraft  10  begins the process by tilting propulsion assembly  22  forward relative to fuselage  18  during the climb. In this configuration, propulsion assembly  22  generates a thrust vector having a forward component  100  that not only urges aircraft  10  to travel in the forward direction but also urges aircraft  10  to rotate about pitch axis  10   b . As the forward airspeed of aircraft  10  increases, collective operation of elevons  30  can be used to enhance the pitch down rotation of aircraft  10 . As seen in  FIG.  2 D , longitudinal axis  10   a  extends out of the horizontal plane H such that aircraft  10  has an inclined flight attitude of about forty-five degrees pitch down. As illustrated, elevons  30  are tilted aftward relative to tail assemblies  26  (see  FIG.  8 C ) and play a progressively larger role in the pitch control of aircraft  10  as the forward speed and inclined flight attitude increase to an aerodynamic flight condition. At the same time, the tilt of propulsion assembly  22  relative to fuselage  18  is preferably being reduced. 
     As best seen in  FIG.  2 E , aircraft  10  has completed the transition to the biplane orientation with forward rotor assembly  24   a  and aft rotor assembly  24   b  counter-rotating in generally parallel vertical planes. In the biplane orientation, wing  14  is above fuselage  18  and wing  16  is below fuselage  18 . By convention, longitudinal axis  10   a  has been reset to be in the horizontal plane H, which also includes lateral axis  10   b , such that aircraft  10  has a level flight attitude in the biplane orientation. As forward flight with wing-borne lift requires significantly less power than VTOL flight with thrust-borne lift, the operating speed and/or collective pitch of forward rotor assembly  24   a  and aft rotor assembly  24   b  may be reduced. In the biplane orientation, the independent control provided by flight control system  20  over elevons  30  and yaw vanes  34   a ,  34   b  provides pitch, roll and yaw authority for aircraft  10  which may be enhanced or complemented with thrust vectoring of propulsion assembly  22 . 
     For example, collective operations of elevons  30  provide pitch authority for aircraft  10  to control, maintain or change the angle of attack of wings  14 ,  16  during forward flight. As best seen in  FIG.  5 A , when each of elevons  30  is tilted forward (see  FIG.  8 B ), the airflow across elevons  30  creates pitch moments having a downward component on elevons  30 , as indicted by arrows  102   a ,  102   b ,  102   c ,  102   d . Pitch moments  102   a ,  102   b ,  102   c ,  102   d  urge aircraft  10  to rotate about pitch axis  10   b , increasing the angle of attack of wings  14 ,  16  and tending to cause aircraft  10  to climb. Similarly, as best seen in  FIG.  5 B , when each of elevons  30  is tilted aftward (see  FIG.  8 C ), the airflow across elevons  30  creates pitch moments having an upward component on elevons  30 , as indicted by arrows  104   a ,  104   b ,  104   c ,  104   d . Pitch moments  104   a ,  104   b ,  104   c ,  104   d  urge aircraft  10  to rotate about pitch axis  10   b , decreasing the angle of attack of wings  14 ,  16  and tending to cause aircraft  10  to descend. 
     As another example, differential operations of elevons  30  provide roll authority for aircraft  10  to control, maintain or change the roll angle of aircraft  10  during forward flight. As best seen in  FIG.  5 C , when elevons  30   b ,  30   d  are tilted forward (see  FIG.  8 B ) and elevons  30   a ,  30   c  are tilted aftward (see  FIG.  8 C ), the airflow across elevons  30  creates roll moments acting generally perpendicularly to elevons  30 , as indicted by arrows  112   a ,  112   b ,  112   c ,  112   d . Roll moments  112   a ,  112   b ,  112   c ,  112   d  urge aircraft  10  to rotate about roll axis  10   a  in the roll left direction, as indicated by arrow  112   e . Similarly, best seen in  FIG.  5 D , when elevons  30   a ,  30   c  are tilted forward (see  FIG.  8 B ) and elevons  30   b ,  30   d  are tilted aftward (see  FIG.  8 C ), the airflow across elevons  30  creates roll moments acting generally perpendicularly to elevons  30 , as indicted by arrows  114   a ,  114   b ,  114   c ,  114   d . Roll moments  114   a ,  114   b ,  114   c ,  114   d  urge aircraft  10  to rotate about roll axis  10   a  in the roll right direction, as indicated by arrow  114   e.    
     The operation of elevons  30  is best seen in  FIGS.  8 A- 8 C  in which a generic tail assembly  26  is depicted. Tail assembly  26  includes a fixed vertical stabilizer  28  and a tiltable elevon  30  proximate the distal end of tail assembly  26 . In the illustrated embodiment, an elevon actuator  106  is configured to tilt elevon  30  relative to tail assembly  26  about elevon axis  108  via linkage  110  responsive to commands from flight control system  20 . When elevon actuator  106  shifts linkage  110  forward, elevon  30  is tilted forward relative to tail assembly  26 , as best seen in  FIG.  8 B . Collectively tilting each elevon  30  of aircraft  10  forward in this manner creates pitch moments  102   a ,  102   b ,  102   c ,  102   d  described above with reference to  FIG.  5 A . Likewise, when elevon actuator  106  shifts linkage  110  aftward, elevon  30  is tilted aftward relative to tail assembly  26 , as best seen in  FIG.  8 C . Collectively tilting each elevon  30  of aircraft  10  aftward in this manner creates pitch moments  104   a ,  104   b ,  104   c ,  104   d  described above with reference to  FIG.  5 B . As described herein, elevons  30  may be differentially operated wherein some of elevons  30  are tilted forward and some of elevons  30  are tilted aftward creating, for example, roll moments  112   a ,  112   b ,  112   c ,  112   d  described above with reference to  FIG.  5 C  or roll moments  114   a ,  114   b ,  114   c ,  114   d  described above with reference to  FIG.  5 D . 
     It is noted that the use of the distributed array of elevons  30  operated by flight control system  20  provides unique advantages for the safety and reliability of aircraft  10  during flight. For example, in the event that flight control system  20  detects a fault with one of the elevons  30 , flight control system  20  is operable to perform corrective action responsive to the detected fault at a distributed elevon level or at a coordinated distributed elevon and propulsion assembly level. As an example and referring again to  FIGS.  5 A- 5 D , flight control system  20  has detects a fault in elevon  30   b  based upon information received from one or more sensors or based upon aircraft response to prior actuation commands. As a first step, flight control system  20  shuts down further operation of elevon  30   b , preferably in a neutral position as represented in  FIG.  8 A . Flight control system  20  now determines what other corrective measures should be implemented based upon the desired maneuvers to be performed by aircraft  10 . For example, flight control system  20  may determine that certain operational changes are appropriate, such as selective use or nonuse of the laterally opposed elevon  30   a  on upper wing  14 , the longitudinally opposed elevon  30   d  on lower wing  16  and/or the diametrically opposed elevon  30   c  on lower wing  16 . In addition to corrective action at the distributed elevon level, flight control system  20  can augment such operations by performing corrective actions with propulsion assembly  22 . 
     For example, to achieve the pitch up maneuver depicted in  FIG.  5 A  during an elevon  30   b  fault, flight control system  20  is configured to leave the laterally opposed elevon  30   a  in the neutral position of  FIG.  8 A  and to actuate elevons  30   c ,  30   d  to the forward tilt configuration of  FIG.  8 B  creating pitch control moments  102   c ,  102   d  that urge aircraft  10  to rotate in the pitch up direction about pitch axis  10   b . In addition, flight control system  20  is configured to coordinate this distributed elevon operation with the upward tilting of propulsion assembly  22  to generate a thrust vector having an upward component, which also tends to urge aircraft  10  to rotate in the pitch up direction about pitch axis  10   b . Similarly, to achieve the pitch down maneuver depicted in  FIG.  5 B  during an elevon  30   b  fault, flight control system  20  is configured to leave the laterally opposed elevon  30   a  in the neutral position of  FIG.  8 A  and to actuate elevons  30   c ,  30   d  in the aftward tilt configuration of  FIG.  8 C  creating pitch control moments  104   c ,  104   d  that urge aircraft  10  to rotate in the pitch down direction about pitch axis  10   b . In addition, flight control system  20  is configured to coordinate this distributed elevon operation with the downward tilting of propulsion assembly  22  to generate a thrust vector having a downward component, which also tends to urge aircraft  10  to rotate in the pitch down direction about pitch axis  10   b.    
     As another example, to achieve the roll left maneuver depicted in  FIG.  5 C  during an elevon  30   b  fault, flight control system  20  is configured to leave the diametrically opposed elevon  30   c  in the neutral position of  FIG.  8 A , to actuate elevon  30   a  to the aftward tilt configuration of  FIG.  8 C  and to actuate elevon  30   d  to the forward tilt configuration of  FIG.  8 B  creating roll control moments  112   a ,  112   d  that urge aircraft  10  to rotate in the roll left direction about roll axis  10   a . In addition, flight control system  20  is configured to coordinate this distributed elevon operation with the creation of a torque imbalance with propulsion assembly  22 . This is achieved by increasing the collective pitch of forward rotor assembly  24   a , decreasing the collective pitch of aft rotor assembly  24   b  or both and/or by increasing the rotor speed of forward rotor assembly  24   a , decreasing the rotor speed of aft rotor assembly  24   b  or both, which also tends to urge aircraft  10  to rotate in the roll left direction about roll axis  10   a . Similarly, to achieve the roll right maneuver depicted in  FIG.  5 D  during an elevon  30   b  fault, flight control system  20  is configured to leave the diametrically opposed elevon  30   c  in the neutral position of  FIG.  8 A , to actuate elevon  30   a  to the forward tilt configuration of  FIG.  8 B  and to actuate elevon  30   d  to the aftward tilt configuration of  FIG.  8 C  creating roll control moments  114   a ,  114   d  that urge aircraft  10  to rotate in the roll right direction about roll axis  10   a . In addition, flight control system  20  is configured to coordinate this distributed elevon operation with the creation of a torque imbalance with propulsion assembly  22 . This is achieved by decreasing the collective pitch of forward rotor assembly  24   a , increasing the collective pitch of aft rotor assembly  24   b  or both and/or by decreasing the rotor speed of forward rotor assembly  24   a , increasing the rotor speed of aft rotor assembly  24   b  or both, which also tends to urge aircraft  10  to rotate in the roll right direction about roll axis  10   a.    
     It is noted that if elevon  30   b  is not shut down in the neutral position as represented in  FIG.  8 A  but instead becomes frozen in an active position such as that of  FIG.  8 B or  8 C , flight control system  20  is configured to take corrective action to overcome this elevon fault at the distributed elevon level. For example, as best seen in  FIG.  5 E , if elevon  30   b  becomes frozen in the tilt forward position depicted in  FIG.  8 B  generating pitch and roll moment  116   b , flight control system  20  is configured to actuate the longitudinally opposed elevon  30   d  to the tilt aftward position depicted in  FIG.  8 C  to generate an opposing pitch and roll moment  116   d . Similarly, as best seen in  FIG.  5 F , if elevon  30   b  becomes frozen in the tilt aftward position depicted in  FIG.  8 C  generating pitch and roll moment  118   b , flight control system  20  is configured to actuate the longitudinally opposed elevon  30   d  to the tilt forward position depicted in  FIG.  8 B  to generate an opposing pitch and roll moment  118   d . In either of these scenarios, flight control system  20  is configured to coordinate this distributed elevon operation with the upward or downward tilting of propulsion assembly  22  to generate a thrust vector having the desired component, to urge aircraft  10  to rotate in the desired direction about pitch axis  10   b  for pitch authority. Likewise, flight control system  20  is configured to coordinate this distributed elevon operation with the creation of a torque imbalance with propulsion assembly  22  by changing the collective pitch and/or rotor speed of forward rotor assembly  24   a  and/or aft rotor assembly  24   b  as required to urge aircraft  10  to rotate in the desired roll direction about roll axis  10   a.    
     As discussed herein, the distributed array of elevons  30  operated by flight control system  20  provides numerous and redundant paths to maintain the airworthiness of aircraft  10 , even when a fault occurs within the distributed array of elevons  30 . In addition to taking corrective action at the distributed elevon level or at the coordinated distributed elevon and propulsion assembly level responsive to a detected fault, flight control system  20  is also operable to change the flight plan of aircraft  10  responsive to the detected fault. For example, based upon factors including the extent of the fault or faults, weather conditions, the type and criticality of the mission, the distance from mission goals and the like, flight control system  20  may command aircraft  10  to travel to a predetermined location, to perform an emergency landing or to continue the current mission. 
     Continuing with the sequential flight-operating scenario, aircraft  10  remains in the biplane orientation with forward rotor assembly  24   a  and aft rotor assembly  24   b  generating forward thrust and with upper wing  14  and lower wing  16  generating wing-borne lift in  FIG.  2 E . In the biplane orientation, flight control system  20  independently controls yaw vanes  34   a ,  34   b  to provide yaw authority for aircraft  10 . For example, yaw vanes  34   a ,  34   b  may be operated collectively to create yaw moments response to the airflow around fuselage  18  and across yaw vanes  34   a ,  34   b . As best seen in  FIG.  5 G , when yaw vanes  34   a ,  34   b  are shifted to the left, the airflow across yaw vanes  34   a ,  34   b  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in a yaw left or counterclockwise direction, when viewed from above, as indicated by arrow  120 . Similarly, as best seen in  FIG.  5 H , when yaw vanes  34   a ,  34   b  are shifted to the right, the airflow across yaw vanes  34   a ,  34   b  creates yaw moments about the center of gravity of aircraft  10  that urge aircraft  10  to rotate about vertical axis  10   c  in a yaw right or clockwise direction, when viewed from above, as indicated by arrow  122 . In the illustrated embodiment, collective yaw vane positioning is achieved by operation of yaw vane actuators  94   a ,  94   b  (see  FIG.  7   ) to shift yaw vanes  34   a ,  34   b  in the same direction relative to fuselage  18  responsive to commands from flight control system  20 . 
     It is noted that the use of propulsion assembly  22  provides unique advantages for maintaining the airworthiness of aircraft  10  in a one engine inoperable event during forward flight. In the illustrated embodiment, forward rotor assembly  24   a  and aft rotor assembly  24   b  of coaxial rotors system  24  respectively receive torque and rotational energy from electric motors  22   g ,  22   h  (see  FIG.  6   ). If flight control system  20  detects a fault with electric motor  22   h , for example, flight control system  20  is operable to perform corrective action responsive to the detected fault including feathering the rotor blades of aft rotor assembly  24   b , which is associated with inoperable motor  22   h , and adjusting the collective pitch and/or rotor speed of forward rotor assembly  24   a , which is driven by the operative motor  22   g . As forward flight with wing-borne lift requires significantly less power than VTOL flight with thrust-borne lift, operation of a single rotor assembly coaxial rotors system  24  provides suitable thrust for continued forward flight. Landing aircraft  10  in the one engine inoperable condition is achieved by transitioning aircraft  10  to the VTOL orientation (see  FIGS.  2 E- 2 G ) and then performing an autorotation and flare recovery maneuver in manned missions with onboard pilot flight control. Alternatively, in both manned and unmanned missions with autonomous flight control, landing aircraft  10  in the one engine inoperable condition is achieved by transitioning aircraft  10  to the VTOL orientation (see  FIGS.  2 E- 2 G ) and then deploying a parachute to reduce the descent speed to a landing surface. 
     Continuing with the sequential flight-operating scenario, as aircraft  10  approaches the destination, aircraft  10  may begin the transition from wing-borne lift to thrust-borne lift. As best seen from the progression of  FIGS.  2 E- 2 G , aircraft  10  is operable to pitch up from the biplane orientation to the VTOL orientation to enable, for example, a vertical landing operation. As seen in  FIG.  2 F , longitudinal axis  10   a  extends out of the horizontal plane H such that aircraft  10  has an inclined flight attitude of about forty-five degrees pitch up. This can be achieved as discussed herein by collective operation of elevons  30 , by thrust vectoring of propulsion assembly  22  or a combination thereof. As illustrated, this causes an increase in the angle of attack of wings  14 ,  16  such that aircraft  10  engages in a climb. In  FIG.  2 G , aircraft  10  has completed the transition from the biplane orientation to the VTOL orientation and, by convention, longitudinal axis  10   a  has been reset to be in the horizontal plane H, which also includes lateral axis  10   b  such that aircraft  10  has a level flight attitude in the VTOL orientation. Aircraft  10  may now commence a vertical descent to a landing surface, as best seen in  FIG.  2 H . As discussed above, during such VTOL operations including hover operations throughout the landing sequence, flight control system  20  may utilize thrust vectoring of propulsion assembly  22  to provide translation authority for aircraft  10  and may utilize differential rotor speed control, differential collective pitch control and/or differential yaw vane positioning to provide yaw authority for aircraft  10 . As best seen in  FIG.  2 I , aircraft  10  has landed in a tailsitter orientation at the desired location. 
     Referring now to  FIGS.  3 A- 3 F  in the drawings, another sequential flight-operating scenario of aircraft  10  is depicted. As best seen in  FIG.  3 A , aircraft  10  is in a tailsitter position on a surface such as the ground or the deck of an aircraft carrier. When aircraft  10  is ready for a mission, flight control system  20  commences operations providing flight commands to the various systems of aircraft  10 . Flight control system  20  may be operating responsive to autonomous flight control, remote flight control, onboard pilot flight control or a combination thereof. As best seen in  FIG.  3 B , aircraft  10  has performed a vertical takeoff and is engaged in thrust-borne lift in the VTOL orientation of aircraft  10 . As illustrated, upper rotor assembly  24   a  and lower rotor assembly  24   b  are counter-rotating in generally parallel horizontal planes. As longitudinal axis  10   a  and lateral axis  10   b  are both in a horizontal plane H that is normal to the local vertical in the earth&#39;s reference frame, aircraft  10  has a level flight attitude. 
     During hover, flight control system  20  may utilize speed control and/or collective pitch control of upper rotor assembly  24   a  and lower rotor assembly  24   b  to cause aircraft  10  to climb, descend or maintain a stable hover. Also during hover, flight control system  20  may utilize thrust vectoring of propulsion assembly  22  to provide translation authority for aircraft  10  and may utilize differential rotor speed control, differential collective pitch control and/or differential yaw vane positioning to provide yaw authority for aircraft  10 . As best seen in  FIG.  3 C , aircraft  10  has completed the vertical ascent to a desired elevation and may now begin the transition from thrust-borne lift to wing-borne lift. As best seen from the progression of  FIGS.  3 C- 3 F , aircraft  10  is operable to pitch down from the VTOL orientation toward the biplane orientation to enable high speed and/or long range forward flight. As seen in  FIG.  3 C , aircraft  10  begins the process by tilting propulsion assembly  22  forward relative to fuselage  18  from a stable hover, instead of a climb as described above with reference to  FIG.  2 C . In this configuration, propulsion assembly  22  generates a thrust vector having a forward component  130  that initially causes aircraft  10  to translate in the forward direction. As the forward airspeed of aircraft  10  increases, forward thrust vector component  130  together with collective aftward tilting of elevons  30  (see  FIG.  8 C ) urge aircraft  10  to rotate about pitch axis  10   b  in the pitch down direction. 
     As seen in  FIG.  3 D , longitudinal axis  10   a  extends out of the horizontal plane H such that aircraft  10  has an inclined flight attitude of about thirty degrees pitch down. As seen in  FIG.  3 E , longitudinal axis  10   a  extends out of the horizontal plane H such that aircraft  10  has an inclined flight attitude of about sixty degrees pitch down. As illustrated, elevons  30  are tilted aftward relative to tail assemblies  26  (see  FIG.  8 C ) and play a progressively larger role in the pitch control of aircraft  10  as the forward speed and inclined flight attitude increase. At the same time, the tilt of propulsion assembly  22  relative to fuselage  18  is preferably being reduced. As best seen in  FIG.  3 F , aircraft  10  has completed the transition to the biplane orientation with forward rotor assembly  24   a  and aft rotor assembly  24   b  counter-rotating in generally parallel vertical planes. In the biplane orientation, wing  14  is above fuselage  18  and wing  16  is below fuselage  18 . By convention, longitudinal axis  10   a  has been reset to be in the horizontal plane H, which also includes lateral axis  10   b , such that aircraft  10  has a level flight attitude in the biplane orientation. In the biplane orientation, the independent control provided by flight control system  20  over elevons  30  and yaw vanes  34   a ,  34   b  provides pitch, roll and yaw authority for aircraft  10  which may be enhanced or complemented with thrust vectoring of propulsion assembly  22 . Thus, aircraft  10  is operable to transition from thrust-borne lift to wing-borne lift taking advantage of the airspeed established in a climb, as discussed with reference to  FIGS.  2 A- 2 E  and is also operable to transition from thrust-borne lift to wing-borne lift taking advantage of the unique thrust vectoring capability of propulsion assembly  22  to generate forward speed, as discussed with reference to  FIGS.  3 A- 3 F . 
     Referring next to  FIG.  9 A  in the drawings, a systems diagram of an aircraft  200  is depicted. Aircraft  200  is representative of aircraft  10  discussed herein. Aircraft  200  includes a propulsion assembly  202 , a gimbal assembly  204 , a flight control system  206 , four tail assemblies  208   a ,  208   b ,  208   c ,  208   d  and a yaw assembly  210 . Propulsion assembly  202  includes a counter-rotating coaxial rotor system  212  formed from rotor assembly  212   a  and rotor assembly  212   b . Rotor assembly  212   a  is operably associated with an electric motor  214   a  and one or more controllers, actuators and/or sensors that are generally designated as electronic systems  216   a , which may specifically include an electronic speed controller, a collective pitch actuator, a health monitoring sensor and the like. Similarly, rotor assembly  212   b  is operably associated with an electric motor  214   b  and one or more controllers, actuators and/or sensors that are generally designated as electronic systems  216   b . Propulsion assembly  202  is configured for omnidirectional thrust vectoring. In the illustrated embodiment, propulsion assembly  202  including rotor assembly  212   a , rotor assembly  212   b , electric motor  214   a  and electric motor  214   b  are tilted relative to the fuselage of aircraft  200  by gimbal assembly  204 . 
     Tail assembly  208   a  includes an elevon  218   a  that is operably associated with one or more actuators, sensors and/or batteries that are generally designated as electronic systems  220   a . Tail assembly  208   b  includes an elevon  218   b  that is operably associated with one or more actuators, sensors and/or batteries that are generally designated as electronic systems  220   b . Tail assembly  208   c  includes an elevon  218   c  that is operably associated with one or more actuators, sensors and/or batteries that are generally designated as electronic systems  220   c . Tail assembly  208   d  includes an elevon  218   d  that is operably associated with one or more actuators, sensors and/or batteries that are generally designated as electronic systems  220   d . Yaw assembly  210  includes yaw vane  222   a  and yaw vane  222   b . Yaw vane  222   a  is operably associated with one or more actuators and/or sensors that are generally designated as electronic systems  224   a . Yaw vane  222   b  is operably associated with one or more actuators and/or sensors that are generally designated as electronic systems  224   b.    
     Flight control system  206  is operably associated with propulsion assembly  202 , gimbal assembly  204 , tail assemblies  208   a ,  208   b ,  208   c ,  208   d  and yaw assembly  210 . In particular, flight control system  206  is linked to electronic systems  216   a ,  216   b ,  220   a ,  220   b ,  220   c ,  220   d ,  224   a ,  224   b  by a fly-by-wire communications network depicted as arrows  226 . Flight control system  206  receives sensor data from and sends commands to propulsion assembly  202 , gimbal assembly  204 , tail assemblies  208   a ,  208   b ,  208   c ,  208   d  and yaw assembly  210  as well as other controlled systems to enable flight control system  206  to independently control each such system of aircraft  200 . 
     Referring additionally to  FIG.  9 B  in the drawings, a block diagram depicts a control system  230  operable for use with aircraft  200  or aircraft  10  of the present disclosure. In the illustrated embodiment, system  230  includes two primary computer based subsystems; namely, an aircraft system  232  and a remote system  234 . In the illustrated implementation, remote system  234  includes a programming application  236  and a remote control application  238 . Programming application  236  enables a user to provide a flight plan and mission information to aircraft  200  such that flight control system  206  may engage in autonomous control over aircraft  200 . For example, programming application  236  may communicate with flight control system  206  over a wired or wireless communication channel  240  to provide a flight plan including, for example, a starting point, a trail of waypoints and an ending point such that flight control system  206  may use waypoint navigation during the mission. In addition, programming application  236  may provide one or more tasks to flight control system  206  for aircraft  200  to accomplish during a mission. Following programming, aircraft  200  may operate autonomously responsive to commands generated by flight control system  206 . 
     In the illustrated embodiment, flight control system  206  includes a command module  242  and a monitoring module  244 . It is to be understood by those having ordinary skill in the art that these and other modules executed by flight control system  206  may be implemented in a variety of forms including hardware, software, firmware, special purpose processors and combinations thereof. Flight control system  206  receives input from a variety of sources including internal sources such as sensors  246 , electronic systems  216 ,  220 ,  224 , propulsion assembly  202 , gimbal assembly  204 , tail assemblies  208  and yaw assembly  210  as well as external sources such as remote system  234 , global positioning system satellites or other location positioning systems and the like. For example, as discussed herein, flight control system  206  may receive a flight plan for a mission from remote system  234 . Thereafter, flight control system  206  may be operable to autonomously control all aspects of flight of an aircraft of the present disclosure. 
     For example, during the various operating modes of aircraft  200  including vertical takeoff and landing flight mode, hover flight mode, forward flight mode and transitions therebetween, command module  242  provides commands to electronic systems  216 ,  220 ,  224 . These commands enable independent operation of each of propulsion assembly  202 , gimbal assembly  204 , tail assemblies  208  and yaw assembly  210 . Flight control system  206  also receives feedback from electronic systems  216 ,  220 ,  224 . This feedback is processed by monitoring module  244  that can supply correction data and other information to command module  242  and/or electronic systems  216 ,  220 ,  224 . Sensors  246 , such as an attitude and heading reference system (AHRS) with solid-state or microelectromechanical systems (MEMS) gyroscopes, accelerometers and magnetometers as well as other sensors including positioning sensors, speed sensors, environmental sensors, fuel sensors, temperature sensors, location sensors and the like also provide information to flight control system  206  to further enhance autonomous control capabilities. 
     Some or all of the autonomous control capability of flight control system  206  can be augmented or supplanted by remote flight control from, for example, remote system  234 . While operating remote control application  238 , remote system  234  is configured to display information relating to one or more aircraft of the present disclosure on one or more flight data display devices  248 . Display devices  248  may be configured in any suitable form, including, for example, liquid crystal displays, light emitting diode displays or any suitable type of display. Remote system  234  may also include audio output and input devices such as a microphone, speakers and/or an audio port allowing an operator to communicate with other remote control operators, a base station, an onboard pilot, crew or passengers on aircraft  200 . The display device  248  may also serve as a remote input device  250  if a touch screen display implementation is used, however, other remote input devices, such as a keyboard or joystick, may alternatively be used to allow an operator to provide control commands to an aircraft being operated responsive to remote control. 
     As discussed herein, aircraft  10  may be a manned or unmanned aircraft and may operate in many roles including military, commercial, scientific and recreational roles, to name a few. For example, as best seen in  FIG.  10 A , aircraft  10  may be a logistics support aircraft configured for cargo transportation. In the illustrated implementation, aircraft  10  is depicted as carrying a single package  300  within fuselage  18 . In other implementations, the cargo may be composed of any number of packages or other items that can be carried within fuselage  18 . Preferably, the cargo is fixably coupled within fuselage  18  by suitable means to prevent relative movement therebetween, thus protecting the cargo from damage and maintaining a stable center of mass for aircraft  10 . In the illustrated implementation, the cargo may be insertable into and removable from fuselage  18  via aft door  18   c  to enable sequential cargo pickup, transportation and delivery operations to and from multiple locations. In one example, aircraft  10  may provide package delivery operations from a warehouse to customers. In another example, aircraft  10  may transport weapons or other military hardware to personnel in a military theater. 
     Aircraft  10  may have remote release capabilities in association with cargo transportation. For example, this feature allows aircraft  10  to deliver cargo to a desired location following transportation thereof without the requirement for landing. For example, upon reaching the desired location in a package delivery operation and transitioned from the biplane orientation to the VTOL orientation, flight control system  20  may cause aft door  18   c  to open such that the cargo can be released from aircraft  10 . This feature may also be useful for cargo drop operations to provide food, water, medicine or other critical items to remote regions during humanitarian or disaster relief missions. Alternatively, as best seen in  FIG.  10 B , the delivery or pickup of cargo may be accomplished using a cargo hook module  302  that may include a cargo hoisting device disposed within fuselage  18  that is operable to raise and lower the cargo while aircraft  10  engages in a stable hover or while aircraft  10  rests in a tailsitting position on a surface. As another alternative, cargo hook module  302  may represent a cargo hook on a fixed length sling assembly that is operable to suspend the cargo a desired distance from the aft end of aircraft  10  during pickup, transportation and drop off of the cargo. 
     As best seen in  FIG.  10 C , aircraft  10  may include a turret mounted sensor assembly  304  that operates one or more sensors such as an integrated sensor suite. For example, sensor assembly  304  may include one or more of an infrared sensor such as a forward-looking infrared (FLIR) sensor, a night vision sensor or other optical sensor, a laser sensor, a lidar sensor, a sound sensor, a motion sensor, a high resolution camera, a radar, a multispectral sensor or any other type of sensor. When aircraft  10  is configured with sensor assembly  304 , aircraft  10  may perform a variety of missions including aerial photography, search and rescue missions, inspection of utility lines and pipelines, environment monitoring, border patrol missions, forest fire detection and monitoring, accident investigation and crowd monitoring, to name a few. In addition, aircraft  10  may engage in military operations such as intelligence, surveillance, target acquisition and reconnaissance. Alternatively, as best seen in  FIG.  10 D , aircraft  10  may be configured to engage in attack missions. In the illustrated implementation, aircraft  10  has a weapons array including four under-wing mounted air-to-ground missile systems  306   a ,  306   b ,  306   c ,  306   d  such as Hellfire or JAGM missile systems. In other implementations, the weapons array of aircraft  10  could include air-to-air missile systems, such as AIM-9 Sidewinder missile systems, and/or anti-submarine torpedo systems such as MK50 torpedo systems. 
     Even though aircraft  10  has been depicted and described herein as having particular attributes, it should be understood by those having ordinary skill in the art that an aircraft could have alternate structure without departing from the principles of the present disclosure. For example, aircraft  310  depicted in  FIG.  11 A  shares many common features with aircraft  10  as indicated by the common numbering of common parts. Aircraft  310 , however, has straight wings, only wing  312  being visible, instead of the swept wings  14 ,  16  of aircraft  10 . Similarly, aircraft  320  depicted in  FIG.  11 B  shares many common features with aircraft  10  as indicated by the common numbering of common parts. Aircraft  320 , however, has flat wings  322 ,  324 , instead of dihedral wing  14  and anhedral wing  16  of aircraft  10 . As another example, aircraft  330  depicted in  FIG.  11 C  shares many common features with aircraft  10  as indicated by the common numbering of common parts. Aircraft  330 , however, has a propulsion assembly  332  that includes a counter-rotating coaxial rotor system  334  formed from rotor assembly  334   a  and rotor assembly  334   b , each of which has two rotor blades instead of four rotor blades as in the rotor assemblies of aircraft  10 . Similarly, aircraft  340  depicted in  FIG.  11 D  shares many common features with aircraft  10  as indicated by the common numbering of common parts. Aircraft  340 , however, has a propulsion assembly  342  that includes a counter-rotating coaxial rotor system  344  formed from rotor assembly  344   a  and rotor assembly  344   b , each of which has five rotor blades instead of four rotor blades as in the rotor assemblies of aircraft  10 . 
     The foregoing description of embodiments of the disclosure has been presented for purposes of illustration and description. It is not intended to be exhaustive or to limit the disclosure to the precise form disclosed, and modifications and variations are possible in light of the above teachings or may be acquired from practice of the disclosure. The embodiments were chosen and described in order to explain the principals of the disclosure and its practical application to enable one skilled in the art to utilize the disclosure in various embodiments and with various modifications as are suited to the particular use contemplated. Other substitutions, modifications, changes and omissions may be made in the design, operating conditions and arrangement of the embodiments without departing from the scope of the present disclosure. Such modifications and combinations of the illustrative embodiments as well as other embodiments will be apparent to persons skilled in the art upon reference to the description. It is, therefore, intended that the appended claims encompass any such modifications or embodiments.