Patent Publication Number: US-2022219814-A1

Title: Rotary-wing, hover-capable aircraft and methods

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims benefit of U.S. provisional patent application Ser. No. 62/852,906 filed May 24, 2019, and entitled “Air Launched Hover-Capable Rotary-Wing Aircraft,” which is hereby incorporated herein by reference in its entirety. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     Not applicable. 
     BACKGROUND 
     Hover-capable, rotary-wing unmanned aircraft, including rotary wing micro air vehicles (MAVs), which rely on the rotation of one or more propellers thereof may have a variety of applications in intelligence, surveillance, reconnaissance (ISR), and search and rescue missions. Hover-capable, rotary-wing aircraft may be electrically powered via one or more batteries carried by the aircraft. In at least some applications, hover-capable, rotary-wing aircraft may have greater power requirements than fixed-wing, non-hover-capable aircraft of a similar size. The power requirements of hover-capable, rotary-wing aircraft may, in at least some applications, may limit the endurance, operating altitude, and/or operating range of the hover-capable, rotary-wing aircraft, thereby limiting the effectiveness of the aircraft. 
     BRIEF SUMMARY OF THE DISCLOSURE 
     An embodiment of a projectile-launched aircraft system comprises a projectile launcher comprising a triggering mechanism, a rotary-wing, hover-capable aircraft comprising a rotor assembly that comprises at least one rotor blade, wherein the rotor blade comprises a stowed configuration and a deployed configuration that is circumferentially spaced from the stowed configuration about a pivot axis, wherein, upon actuation of the triggering mechanism, the projectile launcher is configured to launch the aircraft along a flightpath. In some embodiments, the projectile launcher comprises a barrel configured to receive the aircraft and a cartridge comprising a propellant, and wherein the triggering mechanism is configured to initiate the propellant to launch the aircraft from the barrel. In some embodiments, the flightpath comprises at least one of a vertical flightpath and a ballistic flightpath. In certain embodiments, the aircraft comprises a motor configured to rotate the rotor blade and a control system configured to operate the motor to hover the aircraft at a deployment location that is spaced from the projectile launcher. In certain embodiments, the aircraft comprises an airframe comprising an outer surface comprising at least one first recess formed therein, wherein the first rotor blade is at least partially received in the first recess of the airframe when in the stowed configuration. In some embodiments, the aircraft comprises at least one second rotor blade that is spaced along a longitudinal axis of the aircraft from the first rotor blade, wherein the outer surface of the airframe comprises at least one second recess formed therein, and wherein the second rotor blade comprises a stowed configuration and a deployed configuration that is circumferentially spaced from the stowed configuration about a second pivot axis, and wherein the second rotor blade is at least partially received in the second recess of the airframe when in the stowed configuration. In some embodiments, the aircraft comprises a first gimbal pivotably coupled to the airframe whereby the first gimbal is permitted to pivot relative to the airframe about a first axis, wherein the first rotor blade is coupled to the first gimbal and is permitted to pivot about the first axis relative to the airframe. In certain embodiments, the aircraft comprises a second gimbal pivotably coupled to the airframe whereby the second gimbal is permitted to pivot relative to the airframe about a second axis that extends orthogonally to the first axis, wherein at least one of the first rotor assembly and the second rotor blade is coupled to the second gimbal and is permitted to pivot about the second axis relative to the airframe. In some embodiments, the aircraft comprises a motor assembly configured to rotate the first rotor blade and the second rotor blade, a first servo configured to adjust a position of the first gimbal about the first axis, a second servo configured to adjust a position of the second gimbal about the second axis, and a control system configured to operate the first servo to control a pitch of the aircraft, operate the second servo to control a roll of the aircraft, and to operate the motor assembly to control a yaw of the aircraft. 
     An embodiment of a rotary-wing, hover-capable aircraft comprises an airframe comprising an outer surface that comprises at least one first recess formed therein, a first rotor assembly rotatably coupled to the airframe and comprising at least one first rotor blade, wherein the first rotor blade comprises a stowed configuration and a deployed configuration that is circumferentially spaced from the stowed configuration about a first pivot axis, and wherein the first rotor blade is at least partially received in the first recess of the airframe when in the stowed configuration. In some embodiments, the aircraft further comprises a second rotor assembly rotatably coupled to the airframe and comprising at least one second rotor blade, wherein the second rotor assembly is spaced along a longitudinal axis of the aircraft from the first rotor assembly, wherein the outer surface of the airframe comprises at least one second recess formed therein, and wherein the second rotor blade comprises a stowed configuration and a deployed configuration that is circumferentially spaced from the stowed configuration about a second pivot axis, and wherein the second rotor blade is at least partially received in the second recess of the airframe when in the stowed configuration. In some embodiments, the aircraft further comprises a first motor configured to rotate the first rotor blade in a first rotational direction, and a second motor configured to rotate the second rotor blade in a second rotational direction opposite the first rotational direction. In certain embodiments, the aircraft further comprises a first gimbal pivotably coupled to the airframe whereby the first gimbal is permitted to pivot relative to the airframe about a first axis, wherein the first rotor assembly is coupled to the first gimbal and is permitted to pivot about the first axis relative to the airframe. In certain embodiments, the aircraft further comprises a second gimbal pivotably coupled to the airframe whereby the second gimbal is permitted to pivot relative to the airframe about a second axis that extends orthogonally to the first axis, wherein at least one of the first rotor assembly and the second rotor assembly is coupled to the second gimbal and is permitted to pivot about the second axis relative to the airframe. In some embodiments, the second gimbal is positioned radially within the first gimbal and is configured to pivot about both the first axis and the second axis relative to the airframe, and wherein the first rotor assembly is coupled to the second gimbal. In some embodiments, the aircraft further comprises a motor assembly configured to rotate the first rotor blade and the second rotor blade, a first servo configured to adjust a position of the first gimbal about the first axis, a second servo configured to adjust a position of the second gimbal about the second axis, and a control system configured to operate the first servo to control a pitch of the aircraft, operate the second servo to control a roll of the aircraft, and to operate the motor assembly to control a yaw of the aircraft. In certain embodiments, the aircraft comprises a first rotor assembly that comprises a plurality of circumferentially spaced first rotor blades and a rotor hub centrally positioned between the plurality of first rotor blades, and wherein a radially inner end of each first rotor blade couples to the rotor hub at one of a plurality of hinges. 
     An embodiment of a method for directing a rotary-wing, hover-capable aircraft along a flightpath comprises (a) launching the aircraft from a projectile launcher positioned at a launch location, (b) actuating a rotor blade of the aircraft from a stowed configuration and a deployed configuration that is circumferentially spaced from the stowed configuration about a pivot axis, and (c) hovering the aircraft at a deployment location that is spaced from the launch location. In some embodiments, (b) comprises operating a motor assembly of the aircraft to rotate the rotor blade about a rotational axis. In some embodiments, the first rotor blade is received in a recess formed in an outer surface of an airframe of the aircraft when the first rotor blade is in the stowed configuration. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       For a detailed description of exemplary embodiments of the disclosure, reference will now be made to the accompanying drawings in which: 
         FIG. 1  is a schematic of a projectile-launched aircraft system according to some embodiments; 
         FIGS. 2, 3  are side views of a rotary-wing, hover-capable aircraft of the system of  FIG. 1  according to some embodiments; 
         FIGS. 4, 5  are perspective views of the aircraft of  FIGS. 2, 3 ; 
         FIG. 6  is a side view of the aircraft of  FIGS. 2, 3  opposite the side view shown in  FIG. 2 ; 
         FIG. 7  is a top view of the aircraft of  FIGS. 2, 3 ; 
         FIG. 8  is a perspective view of a powertrain, a first rotor assembly, and a second rotor assembly of the aircraft of  FIGS. 2, 3  according to some embodiments; 
         FIG. 9  is a side view of the powertrain, first rotor assembly, and second rotor assembly of  FIG. 8 ; 
         FIGS. 10, 11  are perspective views of a thrust vectoring assembly of the aircraft of  FIGS. 2, 3  according to some embodiments; 
         FIG. 12  is a flowchart of an embodiment of a feedback control mechanism of the aircraft of  FIGS. 2, 3  according to some embodiments; 
         FIGS. 13, 14  are perspective views of other rotary-wing, hover-capable aircraft of the system of  FIG. 1  according to some embodiments. 
     
    
    
     DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENTS 
     The following discussion is directed to various exemplary embodiments. However, one skilled in the art will understand that the examples disclosed herein have broad application, and that the discussion of any embodiment is meant only to be exemplary of that embodiment, and not intended to suggest that the scope of the disclosure, including the claims, is limited to that embodiment. 
     Certain terms are used throughout the following description and claims to refer to particular features or components. As one skilled in the art will appreciate, different persons may refer to the same feature or component by different names. This document does not intend to distinguish between components or features that differ in name but not function. The drawing figures are not necessarily to scale. Certain features and components herein may be shown exaggerated in scale or in somewhat schematic form and some details of conventional elements may not be shown in interest of clarity and conciseness. 
     In the following discussion and in the claims, the terms “including” and “comprising” are used in an open-ended fashion, and thus should be interpreted to mean “including, but not limited to . . . ” Also, the term “couple” or “couples” is intended to mean either an indirect or direct connection. Thus, if a first device couples to a second device, that connection may be through a direct connection, or through an indirect connection via other devices, components, and connections. In addition, as used herein, the terms “axial” and “axially” generally mean along or parallel to a central axis (e.g., central axis of a body or a port), while the terms “radial” and “radially” generally mean perpendicular to the central axis. For instance, an axial distance refers to a distance measured along or parallel to the central axis, and a radial distance means a distance measured perpendicular to the central axis. 
     As described above, in at least some applications, hover-capable, rotary-wing aircraft may have greater power requirements than fixed-wing, non-hover-capable aircraft of a similar size. The requirements of hover-capable, rotary-wing aircraft may, in at least some applications, limit the endurance, operating altitude, and/or operating range of the hover-capable, rotary-wing aircraft. Additionally, the power requirements of the hover-capable, rotary-wing aircraft may require the use of a battery of increased size which may limit the performance of the aircraft and reduce the amount of payload (e.g., sensors and other equipment) which the aircraft may carry. 
     Embodiments disclosed herein include hover-capable, rotary-wing aircraft which may be launched as a projectile from a launcher positioned a first or launcher location to a second or deployment location distal the launcher location. For instance, the deployment location may be at a horizontal distance along the ground relative to the first location as well as at a different altitude than the launcher location. In this manner, the aircraft may utilize the energy imparted to the aircraft from the launcher to arrive at the deployment location without requiring the operation of one or more propellers of the aircraft. In some embodiments, the aircraft may be launched along a parabolic or ballistic trajectory or flightpath by the launcher towards the deployment location, and may only deploy one or more propellers of the aircraft once the aircraft is either within the vicinity of the deployment location or positioned at the deployment location. In other embodiments, the aircraft may be launched vertically upwards by the launcher towards the deployment location (positioned at an altitude above the launcher location), and may only deploy one or more propellers of the aircraft once the aircraft is either within the vicinity of the deployment location or positioned at the deployment location. 
     Referring to  FIG. 1 , an embodiment of a projectile-launched aircraft system  10  is shown in  FIG. 1 . In the embodiment of  FIG. 1 , system  10  generally includes a projectile launcher  12  and an unmanned hover-capable, rotary-wing aircraft  50  launchable from the projectile launcher  12 . As used herein, the term “hover-capable” describes aircraft capable of hovering flight. Projectile launcher  12  may generally comprise a generally cylindrical firing tube or barrel  14 , a cartridge  16 , and a triggering mechanism or trigger  18 . 
     Barrel  14  of projectile launcher  12  is configured to slidably receive aircraft  50  and, in some embodiments, may have an inner diameter of less than 100 millimeters (mm) (e.g., between approximately 40 mm and 60 mm). Cartridge  16  may also be received within barrel  14  of projectile launcher  12  between aircraft  50  and an enclosed end  15  of barrel  14 . Cartridge  16  may comprise a propellant and an ignition device or primer configured to initiate the propellant in response to receiving a firing signal. In some embodiments, the propellant and primer may each be enclosed within an outer case. The propellant may comprise a material configured to rapidly create pressurized gas to launch aircraft  50  from projectile launcher along an airborne parabolic or ballistic trajectory or flightpath (indicated by arrows  30  in  FIG. 1 ). In some embodiments, the propellant of cartridge  16  may comprise a combustible or explosive material. Trigger  18  of projectile launcher may selectably issue a firing signal to cartridge  16  in response to actuation by a user of projectile launcher  12 . For example, upon actuation, trigger  18  may percussively actuate the primer of cartridge  16  to ignite a propellant of cartridge  16 . 
     As described above, projectile launcher  12  is generally configured to convert energy (e.g., chemical energy) stored within cartridge  16  into kinetic energy of aircraft  50 . Although in the embodiment shown in  FIG. 1  a cartridge  16  is used as a source of energy that may be converted into kinetic energy of aircraft  50 , in other embodiments, the source of energy which may be converted into kinetic energy of aircraft  50  may vary. For instance, pneumatic, hydraulic, and/or electrical sources of energy may be utilized for launching aircraft  50 . Further, in some embodiments, aircraft  50  may not be launched from a cylindrical tube such as the barrel  14  of projectile launcher  12 . For example, aircraft  50  may be launched along an airborne flightpath (e.g., ballistic flightpath  30  shown in  FIG. 1 ) using a rail, catapult, or other mechanism. 
     Aircraft  50  of system  10  may generally include a body or airframe  52  and one or more rotors  54  rotatably coupled to airframe  52 . Rotors  54  may each be disposed in a stowed configuration when aircraft  50  is positioned within the barrel  14  of projectile launcher  12 . In the stowed configuration, each rotor  54  may extend substantially parallel a longitudinal axis of aircraft  50  and positioned against an outer surface of airframe  52  in order to minimize a maximum outer diameter of aircraft  50  and thereby permit aircraft  50  to be loaded into barrel  14 . As will be described further herein, rotors  54  may be actuated from the stowed configuration to a deployed configuration. In the deployed configuration, rotors  54  may be rotated by a motor (not shown in  FIG. 1 ) of aircraft  50  to allow aircraft  50  to hover and perform one or more functions at a deployment location  26  distal the projectile launcher  12 . 
     In some embodiments, the barrel  14  of projectile launcher  12  may be positioned at a non-zero, acute angle  22  relative to the ground  2  to achieve the ballistic flightpath  30  of aircraft  50  following the actuation of the trigger  18  of projectile launcher  12 . For example, in some embodiments, projectile launcher  12  may be light enough to manually aim by a user of projectile launcher  12 ; however, the manner in which projectile launcher  12  may be positioned may vary in different applications. In this embodiment, the barrel  14  of projectile launcher  12  is positioned at acute angle  22  relative to the surface  2  to launch aircraft  50  from a first or launch location  24 , along ballistic flightpath  30 , to the second or deployment location  26  distal the launch location  24 . 
     The deployment location  26  at which aircraft  50  is positioned after being launched from projectile launcher  12  may be spaced a horizontal distance  28  along the ground  2  relative to the launch location  24 . Additionally, the deployment location  26  of aircraft  50  may be spaced a vertical distance or altitude  32  relative to the launch location  24 . Thus, the energy transferred to aircraft  50  from projectile launcher  12  may be utilized to position aircraft  50  at a desired horizontal position along surface  2  and at a desired vertical distance or altitude from surface  2 . By utilizing projectile launcher  12  as an energy source for displacing aircraft  50  from the launch location  24  to the deployment location  26 , energy stored within aircraft  50  (e.g., batteries, fuel) need only be utilized once the aircraft  50  is travelling towards or has arrived at the deployment location  26  for performing one or more functions at the deployment location  26  (e.g., surveillance, etc.). In other words, by utilizing projectile launcher  12  to impart kinetic energy to aircraft  50 , the amount of energy stored within aircraft  50  needed for positioning aircraft  50  at the deployment location  26  may be minimized, thereby maximizing the amount of energy stored within aircraft  50  which may be utilized for performing one or more functions at the deployment location  26 . For instance, by utilizing projectile launcher  12  for transporting aircraft  50  to the deployment location  26 , a payload of aircraft  50  and/or the amount of time aircraft  50  may loiter at operating position  26  may be increased. 
     Although in this embodiment barrel  14  of projectile launcher  12  is positioned at acute angle  22  relative to the ground  2  to produce a ballistic flightpath  30  of aircraft  50 , in other embodiments, the position of barrel  14  may be altered to produce different trajectories of aircraft  50 . For example, barrel  14  of projectile launcher  12  may be positioned at a substantially orthogonal or ninety degree angle relative to the ground  2  to launch aircraft  50  along a vertical flightpath to an operating position that is vertically spaced from projectile launcher  12  but it has minimal or zero horizontal spacing along ground  2  relative to projectile launcher  12 . 
     In his embodiment, the flight of aircraft  50  along ballistic flightpath  30  may include one or more stages. Particularly, the ballistic flightpath  30  of aircraft  50  may include an ascent or projectile phase  34 , a descent phase  36  following the projectile phase  34 , and a hover phase  38  following the descent phase  36 . In some embodiments, during the projectile phase  34  of ballistic flightpath  30  the aircraft  50  is initially launched from the barrel of projectile launcher  12  following the actuation of trigger  18 . During the projectile phase  34 , aircraft  50  ascends or travels vertically upwards away from ground  2  along the ballistic flightpath  30  in response to a force applied to the aircraft  50  from the ignition of cartridge  16 . Additionally, in some embodiments, a protective load transfer cap or cover  56  positioned between aircraft  50  and cartridge  16  within the barrel  14  of projectile launcher  12  to protect components of aircraft  50  (e.g., rotors  54 ) from the shock following the ignition of cartridge  16 . Load transfer cap  56  may fall away from aircraft  50  at some point during the projectile phase  34  of ballistic flightpath  30 . 
     Aircraft  50  may reach the apex of the ballistic flightpath  30  at the end of the projectile phase  34  thereof and begin the descent phase  36 . In the descent phase  36  of ballistic flightpath  30 , aircraft  50  may deploy one or more of the rotors  54  of aircraft  50  prior to reaching the hover phase  38  of ballistic flightpath  30 . In some embodiments, a motor of aircraft  50  may rotate one or more of the rotors  54  whereby the centripetal force applied to the one or more rotating rotors  50  may force the rotors radially outwards (relative a longitudinal axis of aircraft  50 ) from the stowed configuration into a deployed configuration. In other embodiments, different mechanisms may be used to deploy one or more rotors  54  of aircraft  50  during the descent phase. For instance, one or more dedicated actuators of aircraft  50  may be used to deploy rotors  54 . 
     Once rotors  54  of aircraft  50  are disposed in the deployed configuration, a rotational speed of each rotor  54  may be increased whereby thrust sufficient for aircraft  50  to hover above ground  2  and allow aircraft  50  to enter the hover phase  38  of ballistic flightpath  30 . In this embodiment, thrust generated by rotors  54  may be used to slow or terminate the descent of aircraft  50  towards the ground  2  such that aircraft  50  may hover and maintain a desired altitude above ground  2 . In some embodiments, aircraft  50  may include a device configured to induce drag (e.g., a parachute, etc.) during the descent phase  36  to assist in terminating the descent of aircraft  50 . Additionally, in some embodiments, aircraft  50  may include a stabilizer configured to assist in stabilizing the trajectory of aircraft  50  over ballistic flightpath  30 . Aircraft  50  is located at operating position  26  upon reaching the hover phase  38  of ballistic flightpath  30  and may comment performing one or more functions at location  26 . For example, aircraft  50  may comprise a payload including one or more sensors for acquiring data and performing ISR operations at the deployment location  26 . 
     During the travel of aircraft  50  along ballistic flightpath  30 , the motor of aircraft  50  may only be operated during the descent and hovering stages  36 ,  38 . Therefore, the motor of aircraft  50  need not be operated during the ascent phase  34  of flightpath  30  and thus aircraft  50  need only rely on the energy imparted to aircraft  50  from projectile launcher  12  to reach the apex or maximum altitude of ballistic flightpath  30 . In this manner, the amount of energy expended by the motor of aircraft  50  prior to entering the hovering stage  38  at the deployment location  26  may be minimized. 
     Referring to  FIGS. 2-7 , an embodiment of an unmanned hover-capable, rotary-wing aircraft  100  is shown in  FIGS. 2-7 . Aircraft  100  may be utilized in projectile-launched aircraft systems similar to the system  10  described above and shown in  FIG. 1 . Thus, aircraft  100  may be launched from a projectile launcher (e.g., projectile launcher  12 ) at a launch position, travel along an airborne flightpath (e.g., a ballistic flightpath, a vertical flightpath, etc.), and arrive at a deployment position distal the launch position to perform one or more functions (e.g., ISR, etc.). 
     In some embodiments, aircraft  100  may have a longitudinal first end  101 , a longitudinal second end  103  opposite first end  101 , a central longitudinal axis  105  and may generally include a support structure or airframe  102 , a powertrain  150 , a pair of counter-rotating rotors or rotor assemblies  200 A,  200 B, a protective load transfer cap or cover  240 , an actuator or thrust vectoring assembly  250 , and a control system  350 . 
     In some embodiments, aircraft  100  is configured to be launched along a ballistic flightpath (e.g., flightpath  30  shown in  FIG. 1 ) from a projectile launcher (e.g., projectile launcher  12  shown in  FIG. 1 ). Particularly, within a launch tube of the projectile launcher, load transfer cap  240  may be positioned between rotor assemblies  200 A,  200 B (each located at the first end  101  of aircraft  100 ) and a cartridge (e.g., cartridge  16  of  FIG. 1 ) of the projectile launcher. Load transfer cap  240  may include a longitudinal first end  242 , a longitudinal second end  244  opposite first end  242 , the first end  242  being positionable adjacent the cartridge when aircraft  100  is loaded in the projectile launcher. Load transfer cap  240  may protect components of aircraft  100  (e.g., rotor assemblies  200 A,  200 B) from shock generated by the initiation of the cartridge of the projectile launcher. Additionally, load transfer cap  240  may also include a fin or stabilizer  246  proximal first end  242  for stabilizing the flight of aircraft  100  following the launch of aircraft  100  from the projectile launcher; however, in other embodiments, load transfer cap  240  may not include stabilizer  246 . 
     In some embodiments, airframe  102  may provide structural support to and anchor the components of aircraft  100  (e.g., rotor assemblies  200 A,  200 B, thrust vectoring assembly  250 , control system  350 , etc.) and may generally include a generally cylindrical body  104  (body  104  is hidden from view in  FIG. 5 ) and a nose  130  coupled to the body  104 . In some embodiments, components of the airframe  102  (e.g., body  104 , nose  130 ) may be fabricated from acrylonitrile butadiene styrene (ABS) plastic; however, in other embodiments, the materials comprising airframe  102  may vary. 
     Body  104  may comprise a central or longitudinal axis  107  and have a longitudinal first end  105  and a longitudinal second end  106  opposite the first end  105 , and a generally cylindrical outer surface  108  extending between ends  105 ,  106 . In some embodiments, outer surface  108  of body  104  may be smoothed or polished to provide a smooth contact surface between body  104  and an inner surface of a barrel of a projectile launcher from which aircraft  100  is launched. One or more components of aircraft  100  (e.g., thrust vectoring assembly  250 , control system  350 , etc.) may be at least partially disposed within a central passage formed within body  104 . In some embodiments, body  104  may include a removable panel to allow access to components of aircraft  100  stored within body  104 . Body  104  may comprise a pair of circumferentially spaced pivot joints or connectors  110  located at the first end  105  thereof. As will be described further herein, at least a portion of the thrust vector assembly  200  may pivotably couple with body  104  via the pivot joints  110 . In this embodiment, pivot joints  110  are spaced approximately 180 degrees apart about central axis  105  of aircraft  100 ; however, in other embodiments, the circumferential spacing of pivot joints  110  may vary. 
     In some embodiments, the outer surface  108  of body  104  may include a pair of first rotor recesses  112  and a pair of second rotor recesses  116 . Particularly, each rotor recess  112 ,  116  is formed within outer surface  108  of body  104  such that the outer diameter of body  104  defined by outer surface  108  is reduced along the portions of outer surface  108  covered by recesses  112 ,  116 . Additionally, each recess  112 ,  116  extends from first end  105  of body  104  to a terminal end  113 ,  117 , respectively, which is spaced from the second end  106  of body  104 . In this embodiment, the first rotor recesses  112  are spaced approximately 180 degrees apart about central axis  105 . Similarly, the second rotor recesses  116  are spaced approximately 180 degrees apart about central axis  105 ; however, in other embodiments, the circumferential spacing of the pair of first rotor recesses  112  and the circumferential spacing between the pair of second rotor recesses  116  may vary. Additionally, each second rotor recess  116  may be circumferentially spaced from each first rotor recess  112 . 
     As will be discussed further herein, each first rotor recess  112  is configured to receive a rotor blade  202 A of the first rotor assembly  200 A while each second rotor recess  116  is configured to receive a rotor blade  202 B of the second rotor assembly  200 B whereby each rotor blade  202 A,  202 B of the rotor assemblies  200 A,  200 B may be positioned substantially flush with body  104  when each rotor assembly  200 A,  200 B is disposed in a stowed configuration (shown in  FIG. 3 ). In some embodiments, an outer diameter extending between an outer surface  206  of each separate pair of rotor blades  202 A,  202 B (e.g., a diameter extending between the pair of rotor blades  202 A,  202 B of first rotor assembly  200 A) is equal to or less than a maximum outer diameter of the outer surface  108  of body  104 . In this manner, rotor recesses  112 ,  116  of body  104  may serve to minimize a maximum outer diameter of aircraft  100  and protect the rotor blades  202 A,  202 B of each rotor assembly  200 A,  200 B, respectively during the launch of aircraft  100  from a projectile launcher (e.g., projectile launcher  12  shown in  FIG. 1 ). 
     Airframe  102  may additionally include a power supply mount  120  coupled between body  104  and nose  130 , where power supply mount  120  is generally configured to provide structural support to a power supply  140  of aircraft  100 . In this embodiment, power supply mount  120  may comprise an annular flange  122  and a rectangular cage or holder  124  extending from flange  122 . The power supply  140  of aircraft  100  may be received within holder  124 . Particularly, relative movement between power supply  140  and airframe  102  may be restricted when power supply  140  is received within holder  124  and power supply mount  120  is coupled to the body  104  and nose  130  of airframe  102 . 
     In this embodiment, power supply  140  of aircraft  100  comprises an electrical battery pack configured to provide electrical power to components of aircraft  100 , including powertrain  150 , thrust vectoring assembly  250 , and control system  350 . In some embodiments, power supply  140  may comprise a lithium polymer battery configured to output approximately 1,000 milliamp hours (mAh) and 1,500 mAh to allow aircraft  100  to hover for periods in excess of ten minutes; however, in other embodiments, the configuration of power supply  140  may vary. For example, in other embodiments, power supply  140  may comprise fuel storing chemical energy for powering the operation of aircraft  100  rather than a battery pack storing electrical energy. 
     The nose  130  of airframe  102  comprises a longitudinal first end  132  and a longitudinal second end  134  opposite the first end  132  and which defines the second end  103  of aircraft  100 . In some embodiments, the flange  122  of the power supply mount  120  may be coupled between the first end  132  of the nose  130  and the second end  106  of the body  104  of airframe  102 . Nose  130  may include a hemispherical outer surface  136  to reduce the drag of aircraft  100  as it travels along a ballistic flightpath; however, the configuration of nose  130  may vary in other embodiments. 
     Referring to  FIGS. 8, 9 , views of the powertrain  150  of aircraft  100  are shown. In some embodiments, powertrain  150  may generally include a counter-rotating motor assembly  152  that includes a first motor  154  and a second motor  156 . Motors  154 ,  156  of motor assembly  152  may be electrically connected with and powered by the power supply  140  of aircraft  100 . As will be described further herein, control system  350  of aircraft  100  may independently control the operation of each motor  154 ,  156 . First motor  154  is coupled to first rotor assembly  200 A via a first or inner driveshaft  158  which extends through both second motor  156  and second rotor assembly  200 B. Second motor  156  is coupled to second rotor assembly  200 B via a second or outer driveshaft (hidden from view in  FIGS. 8, 9 ) which extends annularly about the inner driveshaft  158 . A fastener or nut  160  may secure rotor assemblies  200 A,  200 B to the motors  154 ,  156  of motor assembly  152 . Additionally, motor assembly  150  may comprise an annular mount  162  positioned about an outer surface  153  of motor assembly  152 . As will be described further herein, mount  162  of motor assembly  152  may pivotably couple with motor assembly  152  to permit motor assembly  152  (along with rotor assemblies  200 A,  200 B coupled thereto) to pivot about a plurality of orthogonal axes relative to the airframe  102  of aircraft  100 . 
     In some embodiments, motors  154 ,  156  may each comprise brushless electric motors separated by bearings (not shown in  FIGS. 8, 9 ) which permit rotors of motors  154 ,  156  to rotate in opposite directions. Additionally, in some embodiments, motor assembly  152  may collectively produce approximately between four and ten Newtons (N) of thrust; however, in other embodiments, the configuration and performance of motor assembly, as well as the relative positioning of motors  154 ,  156  and rotor assemblies  200 A,  2006 , may vary. 
     In the configuration described above, first motor  154  is configured to rotate first rotor assembly  200 A in a first rotational direction (indicated by arrow  155  in  FIG. 8 ) about a rotational axis  151  extending centrally through thrust vectoring assembly  250  while second motor  156  is configured to rotate second rotor assembly  200 B about a second rotational direction (indicated by arrow  157  in  FIG. 8 ) about the rotational axis  151 , where the second rotational direction  157  is opposite of the first rotational direction  155 . In other words, motor assembly  152  of aircraft  100  is configured to counter-rotate rotor assemblies  200 A,  200 B coaxially about the rotational axis  151 . Although rotor assemblies  200 A,  200 B counter-rotate, the rotor blades  202 A,  202 B of rotor assemblies  200 A,  200 B, respectively, are configured to provide a unified or singular thrust vector (indicated by arrow  159  in  FIG. 9 ) in response to the counter-rotation of rotor assemblies  200 A,  200 B. 
     Still referring to  FIGS. 8, 9 , in some embodiments, first rotor assembly  200 A may be spaced along central axis  105  of aircraft  100  from second rotor assembly  200 A. Additionally, first rotor assembly  200 A may comprise a pair of first rotor blades  202 A and a first rotor hub  220 A while second rotor assembly  200 B may comprise a pair of second rotor blades  202 B and a second rotor hub  220 B. In the interest of simplicity, first rotor blades  202 A and first rotor hub  220 A of first rotor assembly  200 A are described in detail below. However, second rotor blades  202 B and second rotor hub  200  of second rotor assembly  200 A may be similar in configuration to first rotor blades  202 A except that second rotor blades  202 B are configured for rotation in the second rotational direction  157  while the first rotor blades  202 A are configured for rotation in the first rotational direction  155 . Additionally, the second rotor hub  220 B of second rotor assembly  200 B may be similar in configuration the first rotor hub  220 A of first rotor assembly  200 A. Thus, rotor blades  202 A,  202 B and rotor hubs  220 A,  220 B include features in common and shared features are labeled similarly. 
     Each first rotor blade  202 A comprises a radially inner end or root  203 , a radially outer end  204 , the outer surface  206  extending between ends  203 ,  204 , a leading edge  208  extending between ends  203 ,  204 , and a trailing edge  210  extending between ends  203 ,  204 . In some embodiments, each first rotor blade  202 A may have a maximum outer diameter (when in the deployed configuration) of approximately between 200 mm and 250 mm, a thickness of approximately between 1.0 mm and 2.0 mm, a 75% span angle of approximately between 15.0 degrees and 20 degrees, a twist of approximately between 6.5 degrees and 8.0 degrees, and a solidity of approximately between 0.05 and 0.08. Additionally, in some embodiments, each rotor blade  202 A may be manufactured using a rapid prototyping technique using polylactic acid (PLA) or carbon reinforced fiber polymers (CRFP). However, in other embodiments, the configuration and process of manufacturing of each first rotor blade  202 A (as well as each similarly configured second rotor blade  202 B) may vary. 
     In some embodiments, the first rotor hub  220 A of each rotor assembly  200 A,  200 B comprises a central passage for receiving the inner shaft  158  of motor assembly  152  and a pair of opposed radially outer ends  222 . Each radially outer end  222  of first rotor hub  220 A may comprise a hinge  224  pivotably connected to the radially inner end  203  of one of the first rotor blades  202 A. In this configuration, each first rotor blade  202 A may pivot about a pivot axis  225  (one of which is shown in  FIG. 8 ) that extends orthogonal central axis  105  of aircraft  100  and which is defined by the hinge  224  coupling the first rotor blade  202 A to the first rotor hub  220 A. Particularly, each first rotor blade  202 A may pivot about pivot axis  225  between a stowed configuration (not shown in  FIGS. 8, 9 ) and a deployed configuration (shown in  FIGS. 8, 9 ). 
     In the stowed configuration, a longitudinal axis of each first rotor blade  202 A of the first rotor assembly  200 A may extend along a longitudinal axis which extends substantially parallel with central axis  105  of aircraft  100 . In the deployed configuration, the longitudinal axis of each first rotor blade may extend substantially orthogonal to central axis  105 . In some embodiments, each hinge  224  of first rotor hub  220 A may include a mechanical stop configured to prevent the first rotor blade  202 A attached thereto from pivoting beyond a substantially orthogonal position (relative central axis  105 ) when the rotor blade  202 A is actuated from the stowed configuration to the deployed configuration. Additionally, in some embodiments, each first rotor blade  202 A may pivot approximately ninety degrees about pivot axis  225  when the first rotor blade  202 A pivots between the stowed and deployed configurations; however, in other embodiments, the relative positioning of the stowed and deployed configurations of each first rotor blade  202 A may vary. 
     In some embodiments, each hinge  224  of first rotor hub  220 A is configured to impart enough friction or resistance to pivoting of each first rotor blade  202 A about its respective pivot axis  225  such that first rotor blades  202 A do not flap (e.g., cyclically pivot about its respective pivot axis  225 ) during operation of aircraft  100 . In some embodiments, following the launching of aircraft  100  from a projectile launcher, the friction imparted by each hinge  224  of first rotor hub  220 A may maintain each first rotor blade  202 A in the stowed configuration until first motor  154  of motor assembly  152  is actuated (e.g., during the decent phase of the flightpath of aircraft  100 ) to rotate in the first rotational direction  155 . In some embodiments, the centripetal force applied to each first rotor blade  202 A in response to the rotation of first rotor assembly  200 A in the first rotational direction  155  overcomes the friction imparted by each hinge  224 , forcing each first rotor blade  202 A radially outwards into the deployed configuration. In other embodiments, an actuator may control the actuation of each first rotor blade  202 A between the stowed and deployed configurations. 
     Referring to  FIGS. 7, 10, and 11 , views of the thrust vectoring assembly  250  of aircraft  100  are shown in  FIGS. 10, 11 . Thrust vectoring assembly  250  is generally configured to control an orientation of each rotor assembly  200 A,  200 B relative airframe  102  to thereby selectably orient or control a vector of the thrust produced by rotor assemblies  200 A,  200 B. The hovering flight of aircraft  100  may be at least partially controlled by vectoring the thrust produced by rotor assemblies  200 A,  200 B using thrust vectoring assembly  250 . 
     In some embodiments, thrust vectoring assembly  250  of aircraft  100  generally includes an annular first or our gimbal  252 , an annular second or inner gimbal  270 , a pitch control rod  290 , a roll control rod  300 , a pitch actuator or servo  310 , and a roll actuator or servo  320 . The pair of pivot connectors  110  of airframe  102  may extend through outer gimbal  252  at circumferentially opposed (e.g., spaced 180 degrees apart about central axis  105 ) locations along the perimeter of outer gimbal  252  to pivotably couple the outer gimbal  252  to the body  104  of airframe  102 . Particularly, a roll control axis  254  may extend through and be defined by the position of pivot connectors  110  whereby outer gimbal  252  and inner gimbal  270  may each pivot relative the airframe  102  about the roll control axis  254 . 
     The inner gimbal  270  of thrust vectoring assembly  250  may couple with the mount  162  of motor assembly  152  via one or more fasteners (not shown in  FIGS. 10, 11 ). Additionally, inner gimbal  270  may be pivotably connected to the outer gimbal  252  by a pair of circumferentially spaced pivot connectors  272  extending radially through the inner gimbal  270  and into a radially inner surface  256  of the outer gimbal  252 . In this configuration, pivot connectors  272  may comprise inner pivot connectors  272  while pivot connectors  110  comprise outer pivot connectors  110 . Each inner pivot connector  272  may be spaced approximately ninety degrees from one of the outer pivot connectors  110 , thereby defining a pitch control axis  274  which extends through inner pivot connectors  272 . In some embodiments, pitch control axis  274  extends orthogonal roll control axis  254 . In addition, roll control axis  254  may extend orthogonally to pitch control axis  274 . Depending on the orientation of outer gimbal  252 , axes  254 ,  274  may each be disposed orthogonal the central axis  107  of the body  104  of airframe  102 . 
     The pitch control rod  290  of thrust vectoring assembly  250  may extend from a first ball joint  292  positioned at a longitudinal first end of pitch control rod  290  to a second ball joint  294  positioned at a longitudinal second end of pitch control rod  290 . The first ball joint  292  of pitch control rod  290  is coupled to an actuator arm  312  that is pivotally coupled to pitch servo  310  at a pivot joint  314 . The second ball joint  294  of pitch control rod  290  is pivotably coupled to the inner surface  256  of outer gimbal  252 . Pitch servo  310  is configured to selectably pivot control arm  312  about a pivot axis defined by pivot joint  314  to linearly displace pitch control rod  290  and, via the pivotable connection between ball joint  294  and outer gimbal  252 , rotate outer gimbal  252  and inner gimbal  270  relative airframe  102  in either rotational direction about pitch control axis  254 . 
     The roll control rod  300  of thrust vectoring assembly  250  may similarly extend from a first ball joint  302  positioned at a longitudinal first end of roll control rod  300  to a second ball joint  304  positioned at a longitudinal second end of roll control rod  300 . The first ball joint  302  of roll control rod  300  is coupled to a control arm  322  that is pivotally coupled to roll servo  320  at a pivot joint  324 . The second ball joint  304  of roll control rod  300  is pivotably coupled to an outer surface  276  of inner gimbal  270 . Roll servo  320  is configured to selectably pivot control arm  322  about a pivot axis defined by pivot joint  324  to linearly displace roll control rod  300  and, via the pivotable connection between second ball joint  304  and inner gimbal  270 , rotate inner gimbal  270  relative to outer gimbal  252  and airframe  102  in either rotational direction about roll control axis  274 . 
     Each servo  310 ,  320  of thrust vectoring assembly  250  comprises a mount  316 ,  326 , respectively, for anchoring each servo  310 ,  320  to an inner surface of the body  104  of airframe  102 , thereby restricting relative movement between servos  310 ,  320  and airframe  102 . Additionally, each servo  310 ,  320  of thrust vectoring assembly  250  may be powered by and electrically connected to the power supply  140  of aircraft  100 . As will be described further herein, each servo  310 ,  320  may be independently controlled by the control system  350  of aircraft to selectably control the attitude and trajectory of aircraft  100  once aircraft enters the hover stage at the deployment location (e.g., deployment location  26  shown in  FIG. 1 ). In some embodiments, aircraft  100  may be controlled by mechanisms other than thrust vectoring assembly  250 , such as via a swash-plate for cyclic and/or collective blade-pitch control for one or both rotor assemblies  200 A,  200 B. 
     Referring to  FIGS. 2, 3, 10, and 11 , Control system  350  of aircraft  100  is generally configured to control the operation of the motor assembly  152  and servos  310 ,  320  of aircraft  100  to control the movement of aircraft  100  once aircraft enters the descent and/or hover stages of the flightpath of aircraft  100 . As shown particularly in  FIG. 5 , control system  350  may generally include a controller or control board  352 , a first motor controller  360 , and a second motor controller  364 . 
     Controllers  352 ,  360 , and  364  may comprise a singular controller or control board or may comprise a plurality of controllers or control boards that are coupled to one another. Controllers  352 ,  360 , and  364  may comprise one or more flexible printed circuit boards (PCB) and/or one or more rigid PCBs with flexible or rigid connections therebetween. Controllers  352 ,  360 , and  364  may each comprise at least one processor and associated memory. The one or more processors (e.g., microprocessor, central processing unit (CPU), or collection of such processor devices, etc.) of each controller  352 ,  360 , and  364  may execute machine-readable instructions provided on the memory (e.g., non-transitory machine-readable medium) to provide each controller  352 ,  360 , and  364  with all the functionality described herein. Additionally, the memory of each controller  352 ,  360 , and  364  may comprise volatile storage (e.g., random access memory (RAM)), non-volatile storage (e.g., flash storage, read-only memory (ROM), etc.), or combinations of both volatile and non-volatile storage. Data consumed or produced by the machine-readable instructions of each controller  352 ,  360 , and  364  can also be stored on the memory thereof. As noted above, in some embodiments, each controller  352 ,  360 , and  364  may comprise a collection of controllers and/or control boards that are coupled to one another. As a result, in some embodiments, each controller  352 ,  360 , and  364  may comprise a plurality of the processors, memories, etc. 
     Controllers  352 ,  360 , and  364  may each be powered by and electrically connected to the power supply  140  of aircraft  100 . Controllers  352 ,  360 , and  364  may collectively control the motion of aircraft  100  once aircraft  100  has entered the hovering stage at the deployment location. Particularly, aircraft  100  may include a center of mass (COM)  190 , a roll axis  192  extending from the COM  190  of aircraft  100 , a pitch axis  194  extending from COM  190  orthogonally to the roll axis  192 , and a yaw axis  196  extending from COM  190  orthogonally to both the roll axis  192  and pitch axis  194 . Controller  352  may be electrically connected or otherwise in signal communication with each servo  310 ,  320  of thrust vectoring assembly  250  and may control the pitch and roll of aircraft  100  by selectably operating servos  310 ,  320 . 
     For example, during the hovering stage of the trajectory of aircraft  100 , controller  352  may selectably actuate pitch servo  310  to rotate motor assembly  152  (coupled to inner gimbal  270  of thrust vectoring assembly  250 ) about pitch control axis  274  to vector the thrust produced by rotor assemblies  200 A,  200 B and thereby induce a pitch moment  193  about the pitch axis  192  of aircraft  100 . Similarly, during the hovering stage of the trajectory of aircraft  100 , controller  352  may selectably actuate roll servo  320  to rotate motor assembly  152  roll control axis  254  to vector the thrust produced by rotor assemblies  200 A,  200 B and thereby induce a roll moment  195  about the roll axis  194  of aircraft  100 . Further, controller  352 , acting through motor controllers  360 ,  364 , may independently vary the rotational rate or revolutions per minute (RPM) of each motor  154 ,  156  of motor assembly  152  such that the RPM of first motor  154  differs from the RPM of second motor  156 . Given that rotor assemblies  200 A,  200 B counter-rotate, a yaw moment  197  may be induced about the yaw axis  196  of aircraft  100  in response to the creation of a differential RPM between motors  154 ,  156 . 
     Referring to  FIGS. 2, 3, and 10-12 , a flowchart illustrating a closed-loop feedback control mechanism  370  implemented by control system  350  is shown in  FIG. 12 . In some embodiments, control system  350  may include an autopilot, an electronic wireless transmitter and corresponding receiver for receiving inputs  372  from a pilot of aircraft  100 , a ground station  374  positioned distal aircraft  100  (e.g., at the launch location  24  shown in  FIG. 1 ), and a telemetry module. Aircraft  100  may be equipped with a custom autopilot along with a telemetry module for stability and to transmit data during flight to ground station  374 . The autopilot of control system  350  may utilize a tri-axial accelerometer and a gyroscope to determine the attitude of aircraft  100 , and further, closed-loop feedback control mechanism  370  and pilot inputs  372  (received via the receiver of aircraft  100 ) to stabilize and control the hovering fight of aircraft  100  at the deployment location (e.g., deployment location  26  shown in  FIG. 1 ). Particularly, the attitude of aircraft  100  may be obtained from the measured body-axis angular rates (gyroscope) and the tilt of the gravity vector (accelerometer). These measurements may be filtered and fused to determine the pitch and roll attitude of aircraft  100  during hovering flight. In some embodiments, measurements of the states of aircraft  100  and control inputs (e.g., pilot inputs  372 , etc.) are transmitted from aircraft  100  to the ground station  374 . 
     As shown particularly in  FIG. 12 , an onboard inner loop feedback  376  of feedback control mechanism  370  corresponding to the body states (e.g., body states p, q, r, φ, and θ) of aircraft  100  is provided by controller  352  of control system  350  while an outer loop feedback  378  corresponding to the inertial states (e.g., inertial states x, y, z) of aircraft  100  is provided by the pilot of aircraft  100  via a controls interface as pilot inputs  372 . In some embodiments, the outer loop feedback  378  provided by the pilot may include heave, roll, pitch, and yaw of aircraft  100 . 
     In some embodiments, electronic control mixing provided by controller  352  provides a plurality of control signals (e.g., pulse width modulated (PWM) signals) for controlling components of aircraft  100 . Particularly, a first control signal  380  may be provided to the first motor controller  360  for controlling the RPM of first motor  154 , and a second control signal  382  may be provided to the second motor controller  364  for controlling the RPM of second motor  156 . In addition, a third control signal  384  may be provided to the pitch servo  310  of thrust vectoring assembly  250  for controlling a command position of pitch servo  310 , and a fourth control signal  386  may be provided to the roll servo  320  of thrust vectoring assembly  250  for controlling a command position of roll servo  320 . 
     In some embodiments, control signals  380 - 386  are processed by a proportional-derivative (PD) controller  388  of feedback control mechanism  370 . Particularly, attitude measurements of aircraft  100  (obtained via the accelerometer and gyroscope of aircraft  100 ) may be fed to PD controller  388  to stabilize the pitch and roll of aircraft  100 . Yaw of aircraft  100  may be stabilized using a derivative feedback controller. Following processing by PD controller  388 , inner loop feedback  376  is provided to a junction  390  which receives pilot inputs  372  and gains and trims  375  from ground station  374 . Particularly, the pilot or other operator of aircraft  100  may update feedback gains, change trim points, and record telemetry data from the autopilot of controller  352  via ground station  374 . 
     In other embodiments, the features of feedback control mechanism  370  of control system  350  may vary. For example, in some embodiments, aircraft  100  may fly autonomously during the hovering stage without input from a pilot, eliminating the outer loop feedback  378  provided by pilot controls  372  and/or ground station  374 . Additionally, in some embodiments, a controller other than PD controller  388  may be used, such as a proportional-integral-derivative (PID) controller or other model-based controllers. 
     In some embodiments, aircraft  100  may include additional sensors and other equipment for performing one or more functions (e.g., ISR, etc.) as aircraft  100  loiters at the deployment location. In some embodiments, sensor data may be transmitted to the pilot or other operator of aircraft  100  via ground station  374 , thereby permitting the operator of aircraft  100  to obtain data pertaining to the deployment location. 
     The configuration of rotary-wing aircraft which may be utilized in projectile-launched aircraft systems (e.g., system  10 ) may vary from the configuration of aircraft  100  shown in  FIGS. 2-12 . For example, referring to  FIG. 13 , another embodiment of an unmanned hover-capable, rotary-wing aircraft  400  is shown. Aircraft  400  may include features in common with aircraft  100 , such as the configuration of control system  350 . However, unlike aircraft  100  which includes a thrust vectoring assembly  250  having nested outer and inner gimbals  252 ,  270 , respectively, aircraft  400  may include two separate and distinct gimbals and rotor assemblies positioned at opposite longitudinal ends of aircraft  400 . 
     Particularly, aircraft  400  has a first longitudinal end  401 , a second longitudinal end  403  opposite the first longitudinal end  401 , a central or longitudinal axis  405 , and may generally include an airframe  402 , a first thrust vectoring assembly  420 , a second thrust vectoring assembly  440 , a first rotor assembly  460 , and a second rotor assembly  480 . First thrust vectoring assembly  420  may include a single gimbal  422  pivotable about a first pivot axis  424  extending orthogonal the central axis  405  of aircraft  400 . A first motor  430  of aircraft  400  may be positioned within the first gimbal  424  for rotating the first rotor assembly  460 , the first motor  430  and first rotor assembly  460  each being rotatable about the first pivot axis  424  relative to the airframe  402 . First rotor assembly  460  is positioned at the first longitudinal end  401  of aircraft  400  and includes a pair of rotor blades  462  and a rotor hub  464  located centrally with respect to rotor blades  462 . A radially inner end or root of each rotor blade  462  may be pivotably connected to rotor hub  464  via a pivot joint  466 . 
     The second thrust vectoring assembly  440  of aircraft  400  may include a single gimbal  442  pivotable about a second pivot axis  444  extending orthogonal the central axis  405  of aircraft  400 . Additionally, second pivot axis  444  may extend orthogonal to the first pivot axis  424  of first thrust vectoring assembly  420 . A second motor  450  of aircraft  400  may be positioned within the second gimbal  444  for rotating the second rotor assembly  480 , the second motor  450  and second rotor assembly  480  each being rotatable about the second pivot axis  444  relative to the airframe  402 . Second rotor assembly  480  is positioned at the second longitudinal end  403  of aircraft  400  and includes a pair of rotor blades  482  and a rotor hub  484  located centrally with respect to rotor blades  482 . A radially inner end or root of each rotor blade  482  may be pivotably connected to rotor hub  484  via a pivot joint  486 . 
     Second rotor assembly  480  may counter-rotate relative first rotor assembly  460  but may, when oriented as shown in  FIG. 13 , produce a singular thrust vector. Additionally, airframe  402  of aircraft  400  may include a first pair of recesses  404  for receiving the rotor blades  462  of first rotor assembly  460 , and a second pair of recesses  406  for receiving the rotor blades  482  of second rotor assembly  480  when each rotor assembly is in a stowed configuration. 
     Referring to  FIG. 14 , another embodiment of an unmanned hover-capable, rotary-wing aircraft  500  is shown. Aircraft  500  may include features in common with aircraft  100 , and shared features are labeled similarly. Particularly, aircraft  500  is similar in configuration except that, instead of nose  130 , the airframe  502  of aircraft  500  comprises a tail  504  that includes a longitudinal first end  506 , a longitudinal second end  508  opposite the first end  506 , and a fin or stabilizer  509  proximal second end  508  for stabilizing the flight of aircraft  500  following the launch of aircraft  500  from a projectile launcher (e.g., projectile launcher  12  shown in  FIG. 1 ). In some embodiments, aircraft  500  may be configured to be launched along a vertical flightpath (indicated by arrow  510  in  FIG. 14 ) from the projectile launcher. Particularly, within a launch tube of the projectile launcher, the tail  504  of aircraft  500  may be positioned adjacent a cartridge (e.g., cartridge  16  of  FIG. 1 ) of the projectile launcher. Thus, unlike the loading of aircraft  100  described above, rotor assemblies  200 A,  200 B of aircraft  500  may be positioned opposite the cartridge of the projectile launcher when aircraft  500  is loaded into the projectile launcher prior to being launched along the vertical flightpath  510 . 
     While embodiments of the disclosure have been shown and described, modifications thereof can be made by one skilled in the art without departing from the scope or teachings herein. The embodiments described herein are exemplary only and are not limiting. Many variations and modifications of the systems, apparatus, and processes described herein are possible and are within the scope of the disclosure. For example, the relative dimensions of various parts, the materials from which the various parts are made, and other parameters can be varied. Accordingly, the scope of protection is not limited to the embodiments described herein, but is only limited by the claims that follow, the scope of which shall include all equivalents of the subject matter of the claims. Unless expressly stated otherwise, the steps in a method claim may be performed in any order. The recitation of identifiers such as (a), (b), (c) or (1), (2), (3) before steps in a method claim are not intended to and do not specify a particular order to the steps, but rather are used to simplify subsequent reference to such steps.