Abstract:
A flying vehicle comprises an elongated airfoil having a first end and a second end defining a longitudinal axis, the first and second ends further defining an interior opening extending therebetween and terminating at an outlet port proximal the second end; a solid fuel engine includes a fuel combustion chamber proximal the first end of the airfoil; a first manifold is coupled to the fuel combustion chamber and configured to convey a pressurized gas generated by combustion of a solid fuel; wherein the interior opening conveys the pressurized gas from the first manifold to the outlet port proximal the second end and wherein the pressurized gas is exhausted at the outlet port substantially perpendicular to the longitudinal axis of the elongated airfoil, thereby producing thrust which imparts a rotational motion to the airfoil about a center of mass proximal to the first end, for thereby defining a rotor disk.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application claims priority benefit and is a divisional of co-pending U.S. patent application Ser. No. 12/786,075, filed May 24, 2010, which application is a divisional of U.S. patent application Ser. No. 11/717,361, now issued as U.S. Pat. No. 7,766,274 filed Mar. 13, 2007, which claims priority to and benefit of U.S. Provisional Application Ser. No. 60/781,361, filed Mar. 13, 2006, the contents of all of which are herein incorporated by reference for all purposes. 
    
    
     FIELD OF THE INVENTION 
     This invention relates to unmanned aerial vehicles, such as those used for surveillance, including helicopter-like “maple seed” or “monocopter” flying vehicles. 
     BACKGROUND 
     The field of unmanned aerial vehicles (UAVs) includes attempts during the American civil war to attack targets by means of unmanned balloons carrying explosive charges. These attempts were generally unsuccessful. During WWI, anti-aircraft gunnery target “drones” were controlled by radio. During WWII, the Japanese used the same balloon bomb technique against the continental United States, resulting in a few deaths. Also during WWII, an Allied program “Operation Aphrodite” attacked surface targets with B-17 bombers converted into drones loaded with explosives. Guidance of the drone B-17s included radio control by a remote operator who viewed a television images from television cameras mounted in the aircraft. In the same general time period, other unmanned aerial vehicles included the German V1 weapons, which were generally unguided in that they were not directed at specific targets, although they were controlled in that they were attitude-stabilized and followed a heading before running out of fuel and crashing. The V2 weapon was also stabilized, and somewhat directionally controlled. More recently, unmanned aerial vehicles have included various missiles such as surface-to-air, air-to-surface, and air-to-air missiles, which are often wholly or partially autonomous, especially in the final attack phase. The Tomahawk “cruise” missile is preprogrammed with a course, and follows the course using Global Positioning Satellite (GPS) positioning and comparison of local sensor data with a preprogrammed digital “map.” 
     Unmanned aerial vehicles have more recently been used for tactical surveillance. This type of vehicle includes the Firebee of Vietnam-war vintage, Hunter, and Pioneer. The Predator is currently in use for surveillance and for other uses. The Predator uses a four-cylinder engine, has a wing span of 48 feet, a length of 27 feet, and a takeoff weight of 2250 lb. It can operate at altitudes of up to 25,000 feet, loiter for up to 40 hours, and in its surveillance role can carry a color video camera, a synthetic-aperture radar, and other sensors. In one of its roles, it can carry and launch Hellfire antitank missiles. The Global Hawk follows the Predator, and provides additional capability, such as a range of more than 12,000 nautical miles and altitudes up to 65,000 feet. However, with a wingspan of 116 feet, length of 44 feet, and 26,000 lb. takeoff weight, it is larger than the Predator. This increased size and weight by comparison with the Predator results in a loiter duration of 24 hours. The Fire Scout is a recently developed reconnaissance and surveillance UAV based on a commercial manned helicopter. The Fire Scout is capable of remote control and of autonomous takeoff and landing from ships or prepared landing sites, and can identify and designate targets. Recent work has been directed toward mounting weapons on the Fire Scout. 
     Ground troops at the small unit level cannot directly take advantage of information resulting from surveillance by unmanned aerial vehicles (UAVs) such as Predator and Global Hawk. However, ground troops would benefit from availability of small reconnaissance UAVs which could examine their local area under the control of the troops themselves, to report directly to those troops. UAVs for such use suffer from various problems, including that they are regularly destroyed or lost during operation, can be damaged by physical abuse, dirt or water, at least some of which tend to be omnipresent in a combat situation. They also tend to have limited range or loiter time, and often require special training to operate. The Dragon Eye backpack reconnaissance transportable UAV is less than two feet long, electrically powered with two propellers, and can be hand- or bungee-launched. Its weight is about 5.5 lb. Its range or loiter time is limited by the capacity of the batteries that can be carried. 
     Improved or alternative unmanned aerial vehicles are desired. 
     SUMMARY 
     An apparatus for flight according to an aspect of the invention comprises an airfoil with an attached payload, and propulsion means associated with the airfoil for rotating the airfoil and the attached payload, for thereby defining a rotor disk. The apparatus also comprises physical means for adjusting the lift of the airfoil, and control means coupled to the physical means for causing the lift adjustment of the airfoil to tilt the rotor disk. In one embodiment, the rotating airfoil defines a leading edge and a trailing or lagging edge, and the physical means comprises means for ejecting gas at a location near the lagging edge in at least one plane which does not coincide with the plane of the rotor disk. In a preferred embodiment, the means for ejecting gas includes means for ejecting gas in a generally periodic manner in a plane not coincident with the plane of the rotor disk. In one embodiment, the airfoil of the apparatus is elongated, and defines a distal end remote from the payload, and the propulsion means comprises a solid-fuel powered bypass jet, with an exhaust directed generally perpendicular to an axis of the elongation of the airfoil. In one version, the exhaust is directed generally in the plane of the rotor disk. The transverse location of the exhaust may lie generally between the distal end of the airfoil and the payload, or it may be at the distal end of the airfoil. 
     According to another aspect of the invention, an apparatus for moving a load in a selected direction comprises an airfoil with a fixedly attached payload, and propulsion means associated with the airfoil for rotating the airfoil and the attached load, for thereby rotating the airfoil to define a rotor disk. Physical means are provided for adjusting the lift of the airfoil. Control means are coupled to the physical means for causing the lift adjustment of the airfoil to tilt the rotor disk in a manner which moves the airfoil with the attached load in the selected direction. 
     A flying apparatus according to another aspect of the invention is for moving a load. The apparatus comprises an airfoil with an attached load fixed to the airfoil, and propulsion means associated with the airfoil for rotating the airfoil and the attached load together, for thereby defining a rotor disk. Physical means are provided for adjusting the lift of the airfoil, and control means are coupled for causing the lift adjustment of the airfoil to provide at least one of collective and cyclic control. 
     A flying apparatus according to a further aspect of the invention comprises an airfoil with an attached load adjacent a first end of the airfoil, and a jet lying between the first and second ends of the airfoil for rotating the airfoil and attached load. 
     An apparatus for flight according to an aspect of the invention comprises an airfoil with a payload which is fixed to the airfoil, and propulsion means associated with the airfoil for rotating the airfoil and the payload, thereby defining a rotor disk. The airfoil with payload fixed thereto has no attached payload which rotates at a rate other than the rotation rate of the airfoil. Physical means adjust the lift of the airfoil, and control means are coupled to the physical means for causing the lift adjustment of the airfoil to tilt the rotor disk. 
     An apparatus for flight comprises an airfoil with a fixedly attached body, where the airfoil and fixedly attached body together defining a center of mass. The attached body includes payload attachment means for attaching a payload centered on the center of mass, which payload, when attached, is fixedly attached to the body. A payload is coupled to the payload attachment means, and propulsion means are associated with the airfoil for rotating the airfoil and the fixedly attached body, for thereby defining a rotor disk. Physical means are provided for adjusting the lift of the airfoil. Control means are coupled to the physical means for causing the lift adjustment of the airfoil to tilt the rotor disk, and control means are provided for controlling the payload attachment means for disengaging the body from the payload at selected one of (a) time and (b) location. 
     An apparatus for flight according to an aspect of the invention comprises an elongated airfoil with an attached payload, which airfoil defines a longitudinal axis. Propulsion means are associated with the airfoil for rotating the airfoil and the attached payload, for thereby defining a rotor disk. The propulsion means comprises means for generating gas under pressure and means for releasing the gas under pressure in a direction generally tangent to a radius of the rotor disk and from a location near an end of the airfoil. The propulsion means in one embodiment of this aspect of the invention comprises an ejector driven by a fuel grain, and the fuel grain may generate hot gas which is partially combustible. In another embodiment according to this aspect of the invention the means for generating gas under pressure comprises a fuel grain which, in operation, creates partially combustible hot gas under pressure, and a first ejector into which the partially combustible hot gas under pressure is introduced, for mixing the partially combustible hot gas with atmospheric oxygen, to generate hot combusted gas. A second ejector receives the hot combusted gas, and heats atmospheric gas to generate the gas under pressure. 
     A protective package according to another aspect of the invention is for individually protecting flying vehicles. The protective package comprises a first piece defining a cavity larger in length, width and depth than corresponding dimensions of the flying vehicle. A second piece is provided having dimensions sufficient to occlude the entirety of the cavity. Means are provided for affixing the second piece to the first piece so as to define with the cavity a closed package containing the flying vehicle. In one embodiment, at least one of the first plastic piece and the second piece is transparent. Either the first piece or the second piece, or both, may be of a plastic material. The second piece may be monolithically hinged to the first plastic piece to define a clamshell. 
     A protective package for accommodating a plurality of flying vehicles according to the invention includes a first generally planar piece defining a plurality, equal in number to the number of the plurality of flying vehicles, of individual open cavities. Each of the open cavities is dimensioned to accommodate one of the flying vehicles. The package also includes a second piece dimensioned to occlude the plurality of individual open cavities. The second piece is applied to the first piece to occlude the open cavities and thereby define the plurality of closed cavities. Each of the closed cavities accommodates one of said flying vehicles. The separate cavities may be separated by a perforated line to aid in separation of the blister-packed vehicles. 
     A method for storing flying vehicles according to a further aspect of the invention comprises the step of encapsulating each flying vehicle in shrink-wrap film, and heating the shrink-wrap film to cause the film to shrink about the flying vehicle. 
     Another method for storing flying vehicles according to an aspect of the invention comprises the steps of placing a flying vehicle in each cavity of a sheet defining plural cavities, and applying a single second sheet over the open side of the plural cavities to form sealed chambers, each holding one flying vehicle. 
     According to another aspect of the invention, the monocopter vehicle includes retaining means for carrying a deployable payload object. In one embodiment, the retaining means comprises a wall defining at least part of a chamber accommodating the deployable payload object. In one version, the wall includes a lip which aids in retaining the deployable payload object, and in another, it includes a withdrawable latch. Yet another version comprises a magnet. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1A  is a plan view of a vehicle according to an aspect of the invention, partially cut away to reveal interior details,  FIG. 1B  is first side elevation thereof,  FIG. 1C  is a second side elevation thereof, and  FIG. 1D  illustrates a rotor disk swept by a rotating structure of  FIGS. 1A ,  1 B, and  1 C; 
         FIG. 2  is a simplified representation of an engine which can be used in the vehicle of  FIGS. 1A ,  1 B, and  1 C, with the engine represented laid out in a straight line for ease of comprehension; 
         FIG. 3A  is a simplified plan view of a portion of the vehicle of  FIGS. 1A ,  1 B, and  1 C, illustrating details of the storage of fuel for the engine of  FIG. 2 ,  FIG. 3B  is a simplified cross-sectional view, laid out straight, illustrating details of one possible structure of  FIG. 3A , and  FIG. 3C  is similar to  FIG. 3B , illustrating another possible structure; 
         FIG. 4A  is a simplified logic flow chart or diagram illustrating how the flyer is controlled, and  FIG. 4B  illustrates a flyer with a line-scan camera in an environment including a street scene; 
         FIG. 5A  is a simplified transverse cross-sectional view of an airfoil of the vehicle of  FIGS. 1A ,  1 B, and  1 C, showing a location for bleeding pressurized gas from the propulsion system, for selectively ejecting the pressurized gas at the trailing edge of the airfoil for increasing or decreasing lift,  FIG. 5B  is a plan view at the location of  FIG. 5A  illustrating the use of a plurality of ejection points for the pressurized gas, and  FIGS. 5C and 5D  together illustrate the effect of the jets on the airflow; 
         FIG. 6A  is a simplified perspective or isometric view of an ejectable payload object enclosed in hinged covers, and  FIG. 6B  is a plan view of the structure of  FIG. 6A  after entering the airstream upon deployment; 
         FIG. 7A  illustrates one possible packaging method for flying vehicles, and  FIG. 7B  illustrates the result of the method; 
         FIG. 8A  is a simplified perspective or isometric view of one possible way to mount the deployable payload object of  FIGS. 6A and 6B  to the vehicle of  FIGS. 1A ,  1 B, and  1 C, and  FIG. 8B  is a plan view illustrating the result of the mounting of  FIG. 8A ; 
         FIG. 9  is a simplified perspective or isometric view of the bottom portion of the payload of the vehicle of  FIGS. 1A ,  1 B, and  1 C, together with a deployable payload object, illustrating a possible retaining means; 
         FIG. 10A  is a simplified plan view of the bottom surface of the payload portion of the vehicle of  FIGS. 1A ,  1 B, and  1 C, showing an alternative structure for deployment of the deployable payload object of  FIGS. 6A and 6B ,  FIG. 10B  is a simplified perspective or isometric view thereof, and  FIG. 10C  is an end view of the structure of  FIGS. 10A and 10B ; 
         FIG. 11A  is a simplified side elevation view of a portion of a vehicle similar to that of  FIGS. 10A ,  10 B, and  10 C illustrating retaining of a deployable payload object by means of a shaped memory actuator,  FIG. 11B  is a simplified perspective or isometric view of the deployable object of  FIG. 11A , and  FIG. 11C  is a detail of the arrangement of  FIG. 11A  showing an alternative state; and 
         FIG. 12  is a notional illustration of various states in the deployment of an embodiment of a deployable payload object. 
     
    
    
     DETAILED DESCRIPTION 
     Helicopters are well known in the art. A helicopter includes a body or payload supported by a rotating airfoil or blade system. The rotating blade system traces out, defines or subtends a rotor disk, which may be understood as producing vertical lift due to the action of the rotating blade system. Each rotating blade of the rotating blade system is an airfoil which is controlled to contribute to the lift, thrust and directional control of the rotor disk, which in turn is imparted to the body. The body, and anything it carries, may be viewed as being the payload which the rotor disk supports. The airfoils of a typical helicopter are connected to a hub region driven by the engine, which is located in the body. The torque applied to provide motion of the airfoils of the rotor disk, by Newton&#39;s laws, cause an equal and opposite torque tending to rotate the body in a direction opposite to that of the airfoils. In order to provide a stationary platform, the rotation of the body of the helicopter is normally controlled, as for example by providing a tail rotor which tends to torque the body of the helicopter in a direction opposite to that engendered by the airfoil rotation torque. Alternatively, coaxial, contrarotating propellers or rotor disks can reduce the torquing of the body, and some cargo helicopters use spaced-apart contrarotating rotor blades. Such helicopters with contrarotating rotor blades may not require powerful auxiliary rotation devices to maintain their attitude. 
     In a typical helicopter, the rotor system includes a rotor head, rotor blades, and a control system that drives and controls the pitch angles of the blade. An axis of rotation is an imaginary line that passes through a point on which the body rotates, where the plane of rotation is orthogonal to the axis of rotation. When a helicopter is flying in a particular “forward” direction, the blades of a helicopter rotor disk alternately speed up (“advancing” blade) and slow down (“retreating” blade) relative to the average airflow, which tends to cause the lift provided by each blade to vary with its rotational position, thereby causing a dissymmetry of lift. Control of the lift and directional control provided by the rotor disk tends to tilt the rotor disk (thereby moving its axis of rotation away from the vertical) to impart “horizontal” directional force as well as lift. This control may be provided in part by a “swash plate” which is coupled to the blades to change the angle of incidence of the blades as a function of rotational position relative to the body of the helicopter. The controls are referred to as “collective,” which acts on all of the blades simultaneously, and “cyclic,” which acts differentially on the blades as a function of rotational position. Thus, “collective” control tends to affect the lift of the rotor disk as a whole, and “cyclic” control tends to tilt the rotor disk. The art of aircraft, including helicopters, and of propellers and rotors, is well known. 
     According to an aspect of the invention, an unmanned aerial vehicle (UAV) without a separate fuselage has a payload or body affixed at one end of a single elongated airfoil. The entire airfoil/payload combination rotates so that the airfoil traces out a rotor disk that behaves much as the rotor disk of a conventional helicopter. The rotor disk may be viewed as being the vehicle, and the rotor blade is simply a lifting rotor. Thrust for rotating the rotor in one embodiment is provided by a jet, which may be tip-mounted. In one advantageous embodiment, the thrust is provided by an air-augmented rocket engine thrusting tangentially at a location remote from the payload, and preferably at that tip of the airfoil remote from the payload. In a preferred embodiment, a control system maintains knowledge of its environment, as by a camera, magnetometer or functionally similar means, to produce directional control signals which actuate lift control means in synchronism with the rotational position of the vehicle. Such a control system allows the vehicle to be directed in a preferred or “forward” direction, and also allows control of lift and therefore of ascent and decent. 
       FIGS. 1A ,  1 B, and  1 C are top plan and first and second elevation views, respectively, of a vehicle according to an aspect of the invention. The description herein may include relative placement or orientation words such as “top,” “bottom,” “up,” “down,” “lower,” “upper,” “horizontal,” “vertical,” “above,” “below,” as well as derivative terms such as “horizontally,” “downwardly,” and the like. These and other terms should be understood as to refer to the orientation or position then being described, or illustrated in the drawing(s), and not to the orientation or position of the actual element(s) being described or illustrated. These terms are used for convenience in description and understanding, and do not require that the apparatus be constructed or operated in the described position or orientation. Terms concerning mechanical attachments, couplings, and the like, such as “connected,” “attached,” “mounted,” refer to relationships in which structures are secured or attached to one another either directly or indirectly through intervening structures, as well as both movable and rigid attachments or relationships, unless expressly described otherwise. 
     In  FIGS. 1A ,  1 B, and  1 C, a vehicle  10  includes an elongated airfoil or blade  12  defining a first or root end  12   1  and a second or tip end  12   2 , and also defines a body or payload designated generally as  14 , which is rigidly affixed to first end  12   1  of the airfoil  12 . The combination of the airfoil  12  and the body or payload  14  has a center of mass designated  8 . The center of mass corresponds with the axis or center of rotation  9  of the structure during operation. The structure defined by vehicle  10  has the same general shape as Samara, the winged fruit of trees such as maples. The vehicle is designated Samarai, blending the name of the fruit with the term “samurai” or warrior. The Samarai is a single-wing helicopter provided with power, flight stabilization, and directional control. 
     Multiple-rotor fuselage-less flying vehicles are known. The concept of a fuselage-less single-blade air vehicle has been known per se since at least the 1970s, as exemplified by the “Maple Seed” toy aircraft, conceived and marketed by Ned Allen. U.S. Pat. No. 5,173,069 was issued on Dec. 22, 1992 in the name of Litos et al., disclosing a toy autorotating flyer having the same general configuration as that shown in  FIGS. 1A ,  1 B, and  1 C. A major advantage of such structures is that they are inherently stable aerodynamically. This toy and other like toys are not useful for surveillance or reconnaissance applications because they autorotate only under the influence of gravity, and can basically only “fly” in a downward direction. For surveillance and reconnaissance purposes, a flying vehicle must be able to at least controllably ascend and proceed in a desired direction. It is advantageous if it can also be controlled to descend. The prior art also includes a rocket-powered “monocopter” flying vehicle, a video of which is available at www.apogeerockets.com/monocopter_movie.asp. 
     The preferred engine components for powering the vehicle of  FIGS. 1A ,  1 B, and  1 C are illustrated in  FIG. 2 . The engine falls into the class of air-augmented rockets (AAR), which are well understood in the art. The engine components will be machined or otherwise formed from high-temperature material, such as nickel. The AAR approach allows much of the oxidizer required for combustion to be off-loaded, and supplied by atmospheric oxygen harvested from the air. Thus, the oxidizer mass of the fuel grain(s) can be less than is usual in a solid-fuel rocket engine. The propulsive efficiency is maximized by the use of air injection and exhaust mass flow augmentation and deceleration, and the specific energy should be comparable to that of a liquid-fuel-based system, and much greater than that of batteries, supercapacitors, or fuel cells. 
     In  FIGS. 1A ,  1 B, and  1 C, the vehicle  10  includes an air-augmented solid-fuel rocket engine designated generally as  210 . Engine  210  includes three stages: an initial stage in which the fuel grain is ignited and burned generating hot high-pressure, fuel-rich gas that exhausted from a nozzle into a second stage. The second stage comprises a supersonic ejector in which the hot, fuel-rich gas from the first stage entrains additional air and establishes a normal shock wave across the throat of that ejector which then allows the second stage to burn the fuel-rich gas creating additional pressure and thrust power. The exhaust from the second stage passes into a third stage, a simple subsonic, thrust augmentation ejector that dilutes the hot exhaust with more air cooling it and increasing the mass flow further.  FIG. 2  is a simplified illustration of the rocket engine  210  which powers the vehicle  10  of  FIGS. 1A ,  1 B, and  1 C. It should be understood that the rocket as illustrated in  FIG. 2  is set out along a straight axis  208  for ease of understanding, but fitting it into the flying vehicle  10  of  FIGS. 1A ,  1 B, and  1 C requires that it be somewhat bent to match the curved shape of the leading wing edge  12   le  of airfoil  12 . In  FIG. 2 , the solid propellant or fuel uses gas-generating power technology. The fuel grain is any grain that produces hot gases that are in part combustible, such as Thiokol type TP-H-3433. The solid propellant lies in a combustion chamber  212  of  FIG. 2 , where an electrical squib ( 329  of  FIG. 3B ) begins a slow combustion. The combustion  213  in chamber  212  produces gas which is more than 50% combustible, and where the temperature is above the ignition temperature of all of the combustible components. The hot combustible gas flows through a nozzle  224  which powers a compressor or primary ejector (which is the essence of the second stage). This compressor is powered by the high-velocity (ideally mach 6 or greater) hot combustible gas from nozzle  224 , and entrains an external air flow  228  at an air intake port  229 . The hot combustible gas and the entrained air enter a mixing chamber  230  defined by walls  226 . The gas flow velocity in the mixing chamber  230  exceeds the flame propagation velocity, so the hot combustible gas does not combust in the mixing chamber, but is mixed and compressed by the narrowing cross-section near a normal shock compressor  240 . The entrained air should have about three times the mass, or more, of the hot gas from nozzle  224 . The precise ratio is a design selection that determines the supersonic ejector performance in terms of compression ratio. The hot gas-air mixture passes through a normal shock compressor  240  and enters a secondary combustion chamber or diffuser  252  defined by walls  250 . The products of secondary combustion in chamber  252  are introduced into a secondary ejector  266  defined by walls  262 , entraining a flow  264  of additional air at an air input port  264 , again at a 3:1 mass ratio, which air is heated by the products of secondary combustion to generate additional gas expansion. The resulting expanded gases under pressure (GUP) leave ejector chamber  266  and flow from an orifice  270  to provide thrust along a tangent  12   t  to the rotating tip end  122  of the airfoil or wing  12 . This thrust rotates ( 10 R) the structure  10  of  FIGS. 1A ,  1 B, and  1 C, including the payload  14 , and the rotating structure  10  traces out or subtends a rotor disk  99  about it center of rotation  9 , as illustrated in  FIG. 1D . Each of the ejectors of engine  210  results in approximately a three-to-one dilution of the hot exhaust gas with air, the full effect being a nine-fold increase in the mass flow and a reduction in the exhaust velocity which increases the thrust efficiency. This tends to greatly increase the mission time including time on station by comparison with a simple rocket or single-stage ejector. The two distinct types of ejectors are desirable as is the combustion process between them in order to achieve maximum efficiency and effectiveness. Experience and analysis have shown that no single ejector can offer more than about three-to-one mass flow augmentation efficiently, so if a 9-to-one augmentation efficiency is desired, two ejectors in series are required. Moreover, experience has shown that the mixing associated with a single ejector tends to consume so much of the enthalpy in the flow that secondary combustion between ejector stage stages becomes desirable to restore it if adequate thrust is to be achieved. If combustion between the ejectors is required, the upstream ejector must operate as a pump and develop a pressure gradient in order to prevent back firing from the secondary combustion. The final ejector can operate at constant pressure device as a simple subsonic mixer provided there no substantial backpressure. 
     The solid fuel grain for the rocket motor or jet  210  of  FIG. 2  is a long cylindrical prism coiled into a continuous spiral to conserve space, or formed into concentric circles of decreasing size toward the center, surrounding the center of mass  8  of  FIG. 1A . The spiral-wound or concentric-ring solid fuel grains lie in a circular area  50  as seen in  FIG. 1A , and in regions  50   a  and  50   b  as illustrated in  FIGS. 1B and 1C . 
       FIG. 3A  illustrates details of the propellant or fuel storage of the structure  10  of  FIGS. 1A ,  1 B, and  1 C. In  FIG. 3A , the region in which the fuel grains are stored is designated  50 . Each fuel grain is located within a tube or a hollowed-out groove within a refractory material, a set  310  of three of which are illustrated, namely  310   a ,  310   b , and  310   c . It will be appreciated that more or fewer tubes than three can be used. The tubes of set  310  are coiled around the center of mass  8  of the vehicle. A fuel gas manifold  224  connects the openable ends of the tubes of set  310 , and leads the combustible gas into the first ejector housing  226 . The actual first ejector air intake is designated  229 .  FIG. 3B  illustrates the general shape of a fuel grain tube, designated  310   c  for definiteness, laid out in a straight line. In  FIG. 3B , the metal tube wall is designated  310   cw  and the solid fuel grain is designated  310   cg . The fuel grain is the Thiokol type TP-H-3433, or similar, which when ignited by an igniter illustrated as  329  begins partial combustion at the exposed face, generating hot combustible gas. The hot combustible gas enters manifold  224  for energizing the primary ejector  224 ,  226 ,  230  of  FIG. 2 . Fuel ignition is begun in one of the tubes of set  310  by squibs or semiconductor bridge initiators, known in the art. Such initiators operate quickly, on the order of 100 microseconds (μs), and require only about 3 millijoules (mJ) of energy. Once a grain is burning, it is virtually impossible to stop. Some control of the length of burn is achieved by making the grains short, and since they are positioned in parallel, a fresh one can be ignited just before the end of burn of the previous grain, or, if desired, the current grain can be allowed to burn itself out without igniting another. The burn rate of a solid fuel grain such as  310   cg  of  FIG. 3B  depends upon the pressure at the face. Some control of the pressure can be achieved by a controllable orifice, such as an orifice  222  illustrated by a valve symbol  218 . Any type of microelectromechanical valve  218  can be used to adjust the pressure via electrical control  220 . More control of the pressure may be possible with an arrangement such as that illustrated in  FIG. 3C , in which the fuel grain  310   cg  defines an open axial bore  310   cb  which extends the full length of the grain. The pressure-controlling valve  218  is affixed to the end  216  of tube wall  310   cw  remote from the ignition device  329  and the manifold  224 . This arrangement allows the pressure at the burn face to be adjusted to thereby throttle the rocket/ejector apparatus. 
     In order to prevent cross-ignition of the fuel grains at the manifold  312  of  FIG. 3 , each fuel grain must be insulated sufficiently from its neighboring grains to prevent inadvertent ignition. The refractory material that houses and surrounds the fuel grain must have sufficient insulation value to accomplish this task. Any of the refractory oxides known in the art (e.g., alumina, zirconia, etc) can serve here. Additional cooling passages can be added if desired when using especially hot fuels or fuels with low ignition temperatures. 
     Takeoff of the vehicle  10  of  FIGS. 1A ,  1 B, and  1 C, powered by a rocket or jet engine, may be aided by a launching device which brings the vehicle  10  up to full rotational speed and ignites the engine when full speed has been reached, in order to conserve on fuel, but this is not believed to be mandatory. The provision of such a launching device is believed to be a simple matter easily made by a person skilled in the art. 
       FIG. 1A  illustrates by  40  a possible location for a solid-state microchip radio/television transceiver (transmitter/receiver). Those skilled in the art know that the efficiency of an antenna in transducing signal is dependent, in part, upon its dimensions. Ideally, communications with the vehicle  10  occur at frequencies at which the available antenna dimensions are sufficient to transduce signals with an efficiency which allows reliable communication with the outside world with an amount of electrical power available from an on-board thin-film battery  42 . Moreover, the antenna can be fitted with a backplane or reflector to further improve efficiency. Also, transmission may be pulsed so that a directed beam is emitted from the Samarai only when the antenna is facing the right direction. 
     Block  44  of  FIG. 1A  represents microchip television camera and other imaging sensors and electronics which are powered by the battery  42 . The images and sensed signals may include part of the information for which the reconnaissance is performed, so the image and sensor information from block  44  may be coupled to the transceiver  40  for transmission to the home base. 
     In one embodiment of the vehicle  10  of  FIGS. 1A ,  1 B, and  1 C, the length of the vehicle from the center of rotation  9  to the tip end  12   2  of the airfoil  12  is expected to be 3.5 centimeters. The weight of the vehicle  10  with payload and full load of fuel (2 grams of Thiokol type TP-H-3433 or similar) is expected to be in the region of 10 grams, including a deployable or ejectable 2-gram payload. The spin rate is expected to be 250 Hz, the disk loading 0.25 kg/m 2 , and the endurance not less than 20 minutes. 
     Navigation and proximity sensors are represented in  FIG. 1A  by microchip set  46 . One of the problems associated with a vehicle according to an aspect of the invention is that the entire body and rotor rotate, so that any sensors incorporated into the structure and which are used for navigation and/or control are spinning about the axis of rotation  9  ( FIG. 1B ). According to an aspect of the invention, the rotation rate of the craft is determined by the use of a video or television camera which views the outside world, and recurrently sees the same features of the environment as the vehicle spins. This information is used for estimating or determining the spin speed or rate. A clock in an on-board processor determines the angle of the rotor blade relative to the rotor disk using the spin information derived from the camera. A line-scan television camera produces images representing 360° around the environment of the vehicle, including image portions representing the “forward” direction. Real-time imagery in the forward direction can be generated by “de-spinning” the image. The de-spinning can be accomplished on-board, or can be offloaded to an external controller. The controller in such a case would require knowledge of the spin rate and the rotor blade angle relative to the rotor disk. As an alternative to a line-scan camera, a CCD chip could “snap” a picture at that time at which it faces forward to thereby produce visual signals. In either case, the image information can be used for autonomous control or transmitted to a remote controller for external instructions. Ideally, the vehicle stability control and detailed aspects of the altitude, direction, and attitude control would be controlled autonomously by the vehicle, leaving only higher-level instructions to the external controller. Such higher-level instructions might include heading and altitude, or go-right/go-left instructions, without detailed attitude instructions. Autonomous control would also be advantageous for engine thrust control and propellant usage minimization. 
     The art of remote control and autonomous control of flying vehicles, including helicopters, is well advanced.  FIG. 4A  is a simplified block diagram of a controller responsive to images from a line-scan camera  446  of  FIG. 4B  to determine the spin, and for communicating with an autonomous controller  448  or by radio (or equivalent)  449  with a remote control source  450 .  FIG. 4B  illustrates a flyer such as  10  of  FIG. 1A  carrying a line-scan camera  446  which repeatedly scans a line  447  across a street scene illustrated as  444 . A vertical edge  442  in the scene, or any other vertical edge, can be used as a reference. In general, the “forward” direction is determined by estimating the angle of the airfoil or blade relative to the rotor disk using a precision clock in an on-board processor synchronized with the rotation rate and phase. The arrangement of  FIG. 4A  receives a scene input from the scanning camera  446 , as suggested by block  412 , and produces a signal which represents the angle of the rotor blade relative to an arbitrarily selected “front” direction. A block  414  represents the synchronization of a virtual clock signal with the repeating image, and the identification of one or more reference vertical features of the scent. The stabilized scene information is transmitted to a remote control location, as suggested by block  449 . Block  416  represents the correlation of the data extracted from the camera image with data from other sensors, such as a horizontally-disposed laser range finder or magnetometer. Block  418  represents the controlled actuation of gas valves, as described in more detail in conjunction with  FIGS. 5A ,  5 B,  5 C, and  5 D to implement motion in the selected “forward” direction. The “forward” direction information is provided from block  416  to an autonomous controller illustrated as a block  448 . Remote-control and autonomous systems for control of unmanned helicopter vehicles are known. The principles of such systems can be adapted for use with a vehicle according to an aspect of the invention. When operating autonomously, controller  448  produces signals which are ultimately made available to the control valve subroutine of block  418  to control the attitude and direction of the vehicle. 
     The line-scan image is stabilized by addressing the image information using the tracking clock described in conjunction with  FIG. 4 . A portion of the image is selected as representing the “front” view, and further control uses this information to aid in navigation. A 128-pixel vertical-line scan camera is believed to be sufficient to provide a picture with sufficient resolution to allow an operator to guide the vehicle into tight spaces, while placing a relatively low load on the communication system. Bandwidth compression techniques can be used if desired when transmitting the camera pictures. 
     As so far described, the vehicle is spinning in an aerodynamically stable manner, and the “forward” direction of travel is known. Instructions are generated either autonomously or from an external controller to proceed in some selected direction. For simplicity, an explanation of the forward direction is selected. In order to cause the vehicle to move in the selected direction, the rotor disk must tilt relatively upward at the “rear,” or downward at the “front.” In order for the rotor disk to tilt, the rotating airfoil or blade must exhibit differential lift. More particularly, the lift when the airfoil is at the 180° (rearward) position on the rotor disk must be greater than the lift when the airfoil is at the 0° (forward) position on the rotor disk. Control of the lift of the airfoil is accomplished according to an aspect of the invention by selectively ejecting pressurized gas from the trailing edge of the airfoil in a direction selected to increase or decrease lift, as a function of angular position of the airfoil or blade relative to the rotor disk.  FIG. 5A  is a simplified diagram representing a cross-section of the airfoil or blade of  FIGS. 1A ,  1 B, and  1 C at a location designated  5 - 5  in  FIG. 1A . As illustrated in  FIG. 5A , the housing  262  of chamber  266  opens at this particular cross-section into a tube-fed manifold arrangement  512 ,  514 . Chamber  266  provides relatively cool, pressurized gas by way of manifold  514  to a pair of microelectromechanical valves  521 ,  522 . Valves  521  and  522 , when electrically energized, allow the flow of the pressurized gas to upwardly-directed jet  531  and downwardly-directed jet  532 . Jets  531  and  532  are mounted at or immediately adjacent to the trailing edge  12   te  of the airfoil  12 .  FIG. 5C  is a simplified cross-sectional view similar to  FIG. 5A , illustrating only valve  532  and upwardly-directed jet  531 , during a time at which the valve  532  is closed to prevent the flow of gas to the jet. Under this condition, the streamlines, illustrated as  550 , flow smoothly over the region of jet  531 . The upwardly directed jet  531 , when valve  521  is opened to allow gas to flow, directs the flow of gas in an upwardly skewed plane  598   u  relative to the rotor plane, as illustrated in  FIG. 5A . This creates a separation or steering bubble  560  adjacent jet  531  as illustrated in  FIG. 5D . The separation bubble forces the airflow, as represented by streamlines  550 , up and over the bubble  560 . The effect is equivalent to adding more camber to the airfoil at the location of the jet  531 , thereby increasing the lift at that point. Moving the chordwise lift distribution toward the trailing edge changes the pitching moment balance and causes the airfoil to pitch nose down, thereby reducing angle of attack of the airfoil and its overall lift production. Downwardly-directed jet  532  of  FIG. 5A  directs the gas to the underside of the airfoil and performs the symmetrically opposite function as  531  for the upper surface. In a preferred embodiment, only one valve and jet, say  521  and  531 , is needed—the other being redundant. In order to tilt the rotor disk to direct the vehicle in the forward direction, the lift of the airfoil is increased when at the 180° (rearward) position on the rotor disk, by closing microelectromechanical valve  521  to cut off the supply of gas to jet  531 , thereby reducing the nose-down torque on the airfoil, allowing its angle of attack to increase and its lift production to increase. A half-rotation later, when the airfoil is at the 0° (forward) position on the rotor disk, the process is reversed: valve  521  is open, gas is flowing into the bubble increasing the downwash from the trailing edge and thus the nose down torque, reducing the angle of attack and the lift. The valve or valves is (are) actually operated many times as the airfoil advances around the stations of the compass, so that the size of surface bubble waxes and wanes evenly, or nearly so, during each circuit of the airfoil. Thus, the rotor disk experiences increased lift at its rearmost position and relatively decreased lift at its forward portion. The resulting tilt of the rotor disk causes the vehicle to tend to move in the selected forward direction. 
       FIG. 5B  is a simplified plan view of a portion of a region of the airfoil  12  of  FIG. 5A , partially cut away to reveal interior details. In  FIG. 5B , it can be seen that controllable valve  521  can control the compressed gas to a plurality of upwardly-directed jets, designated  531   a ,  531   b ,  531   c ,  531   d , and  531   e , spaced along the trailing edge  12   te  of the airfoil  12 . If desired, a control valve equivalent to valve  514  can be used to separately control the flow of compressed gas to each of the individual upwardly-directed jets  531   a ,  531   b ,  531   c ,  531   d , and  531   e . It will be understood that there may be a plurality of downwardly-directed jets, controlled by valve  522  in the same manners as described for the upwardly-directed jets. 
     As so far described, the vehicle is adapted to reconnaissance by virtue of its ability to fly into desired regions, and to use sensors to report conditions at the desired locations. The television camera can provide very valuable information to a remotely located operator. The “payload” as so far described has been the body portion which is fixed to the root end of the airfoil, and which carries the various sensors. It may sometimes be advantageous to be able to drop an object, such as a remote sensor, from the airborne vehicle. The conventional name for such an object is also “payload.” In order to avoid confusion, the object dropped is called an “object.” Referring again to  FIG. 1B , an ejectable payload object is illustrated as a disk-like object designated  14   ep , which is attached or affixed to the “bottom” of the payload portion  14  of the vehicle  10 .  FIG. 6A  is a simplified illustration of a payload object  14   ep  in its stowed state, suitable for being transported by the vehicle, and  FIG. 6B  is an illustration of the payload object  14   ep  in a state following deployment or ejection from the vehicle. In the stowed state illustrated in  FIG. 6A , the payload object  14   ep  has the form of a right circular cylinder defined around an axis  608 . The payload object  14   ep  incorporates first and second “drag wings”  610   a  and  610   b , respectively, folded against the sides of the payload object. These wings are hinged so that they can open to extend radially outward from the payload object when the payload object is released from the body of the vehicle. First drag wing  610   a  defines a semiannular side wall  610   asw  and half-circular top wall or surface  610   aus . First half drag wing  610   a  is hinged to the underlying structure by a vertically disposed hinge  610   ah . Similarly, second drag wing  610   b  defines a semiannular side wall  610   bsw  and an upper half-circle top wall or surface  610   bus . Second drag wing  610   b  is also hinged to the underlying structure by a vertically disposed hinge, a portion of which is illustrated as  610   bh . The first and second semiannular side walls are separated from each other by a separation, part of which is designated  620 . 
       FIG. 6B  is a simplified top view illustrates the state of the payload object  14   ep  of  FIG. 6A  after it is deployed or ejected from the vehicle  10  of  FIGS. 1A ,  1 B, and  1 C. Since the vehicle  10  spins in normal flight, the deployable payload object  14   ep  also spins. On release of the payload object from the vehicle body, the drag wings spring open assisted by centrifugal forces and quickly reduce the rotational energy and rotation rate (the rotation energy is proportional to the square of the rotation rate, so the energy drops off faster than the rate. When so opened or deployed, the drag wings act as anti-spin spoilers or air brakes, which tend to reduce the spin. Reduction of spin is desirable to aid in precise placement of the payload object  14   ep  in the target area, by preventing the object from “skittering” and moving uncontrollably energy. This skittering is also reduced by the fact that the drag wings are fashioned from an elastic material, as opposed to or vice a rigid material, so that they absorb shock. On contact with an alightment surface, the drag wings are torn away revealing an adhesive  614  of  FIG. 6B , that helps the payload object remain where it lands. 
     It will be clear that if the anti-spin spoiler or drag wings of the payload object are exposed to the airstream while the payload object is being carried, the drag wings will open spontaneously, which might introduce excessive rotational drag while the payload object is still attached to the airfoil/payload combination.  FIG. 8A  is a simplified perspective or isometric view of a payload  14  portion of a flying vehicle  12 , showing the bottom surface  14   bs  of the payload  14 . A wall arrangement  810  includes an attached pair of partially circular or semicircular walls  810   a ,  810   b , which together define a generally cylindrical cavity  812 , having at least a diameter suitable to accommodating the payload object  14   ep . The function of the walls  810   a ,  810   b  is to prevent wind action from seizing the drag wings  610   a ,  610   b  and opening them during flight. The height or projection of the walls  810   a ,  810   b  can be the same as the height of the right circular cylinder defined by the payload object  14   ep , in which case the drag wings  610   a ,  610   b  are pretty much protected against wind. As an alternative, the height of walls  810   a ,  810   b  can be less than the height of the right circular cylinder defined by payload object  14   ep , in which case the drag wings are partially protected from wind. Being only partially protected, there might be some forces tending to open the drag wings  610   a ,  610   b , but they cannot open in any case because of the presence of the walls  810   a ,  810   b  standing in the path of opening.  FIG. 8B  is a plan view of the lower surface of the structure of  FIG. 8A , showing the walls  810   a ,  810   b . As an alternative, the two walls  810   a ,  810   b  may be joined by additional wall segments so as to define a complete 360° enclosure, thereby providing more protection against wind. 
     The payload object  14   ep  can be held to the underside  14   bs  of the payload of the vehicle  10  by a magnetic device. If the payload object  14   ep  includes a magnetically permeable material, a simple electromagnet mounted on the bottom or lower surface  14   bs  can hold the payload object in place until such time as ejection or deployment is desired.  FIG. 9  is an exploded perspective or isometric view illustrating, in notional form, a controllable electromagnet  910  associated with the bottom or payload-object-supporting side of the payload body  14  of the vehicle  10 .  FIG. 9  also illustrates the payload object  14   ep  with a magnetically permeable portion  812 , which is either inherently part of the payload object, or is added to allow controlled stowage, carriage and deployment. The magnetically permeable portion  812  is centered on that side or surface  14   epus  of the payload object  14   ep  which faces the bottom surface  14   bs  of the payload  14  of the vehicle  10 , where it can interact with electromagnet  910 . When the electromagnet  910  is actuated, the deployable payload object  14   ep  is held to the vehicle  10 , and when the electromagnet  910  is deenergized, the deployable payload object is released or deployed. If the magnetically permeable portion  812  holds some remanent field, deployment might be aided by reversing the magnetic field of electromagnet  910 , thereby tending to repel the deployable payload object  14   ep.    
     As an alternative to the electromagnet described in conjunction with  FIG. 9 , the deployable payload  14   ep  might be mounted and held slightly away from the center of mass of the vehicle  10 , so that some small centrifugal force tends to eject the deployable payload object. A microelectromechanical latch can hold the deployable payload object against the small force, but release it when actuated. This would tend to fling the deployable payload object in a direction related to the rotational position of the vehicle at the time that the microelectromechanical latch releases the deployable payload object  14   ep , which might be advantageous for some scenarios.  FIG. 10A  is a plan view of the bottom surface  14   bs  of the payload or body portion  14  of a vehicle  10  which illustrates one possible way to accomplish the desired result. In  FIG. 10A , the bottom surface  14   bs  of the payload portion  14  bears a retaining and wind-protecting wall  1010 , which includes a circular portion centered on a point  1012 , somewhat displaced in the direction of arrow  1014  from the center of mass  9  of the vehicle  10 . The right-circular-cylindrical deployable payload object  14   ep  lies against the circular portion of wall  1010 , and slightly offset from the center of mass  9 . Rotation of the vehicle  10  tends to cause a force tending to move the deployable payload object  14   ep  in the direction of arrow  1014 . Movement of the deployable payload object  14   ep  in the direction of arrow  1014  from the illustrated rest position is prevented by one or more microelectromechanical trip devices designated together as  1016 . These devices when holding the deployable payload object project from the bottom surface  14   bs  of the payload portion  14  of the vehicle, preventing the deployable object from moving in the direction of arrow  1014 . When tripped, the trip device(s)  1016  are either ejected or withdrawn under surface  14   bs , in either case releasing the deployable payload object  14   ep  to move in the direction of arrow  1014  under the impetus of inertial forces, thereby accomplishing the deployment.  FIG. 10B  is a perspective or isometric view of the arrangement of  FIG. 10A , showing that the side walls  1010  affixed to the lower surface of the flyer  10  may have retaining lips  1050  which overhang a portion of the deployable payload object  14   ep  to retain it in place, allowing egress only by way of the “opening”  1040 . As illustrated in  FIG. 10B , the retaining trip devices are three in number.  FIG. 10C  is an end view of the structure of  FIG. 10B  looking in the direction of section lines  10   c - 10   c.    
       FIGS. 11A ,  11 B, and  11 C illustrate another arrangement for stowing the deployable payload object  14   ep . In  FIGS. 11A and 11B , the deployable payload object  14   ep  includes at least one retaining tab or tang  1114  which fits under an overhanging lip portion  1114 L. The deployable payload object also defines a shaped depression  1014   d  (not visible in  FIG. 11B ). When the deployable payload object  14   ep  of  FIG. 11A  is mounted to the air vehicle body  14 , the retaining tang can be fitted under lip  1114 L as illustrated in  FIG. 11A , and the other side of the payload object  14   ep  can be fitted into a recess  14   r  with shaped depression  1114   d  aligned with a retaining latch  1120 . The retaining latch  1120  can be a shape memory actuator which can be energized by piezoelectric effects, magneto-resistive effects, or thermal heating effects, as known in the art. The retaining latch  1120  assumes the locked position illustrated in  FIG. 11A , and can be energized to the unlocked position illustrated in  FIG. 11C , which releases the latch  1120  from the shaped depression  1014   d  to thereby release the deployable payload object  14   ep  of  FIG. 11A  from the body  14 . A spring or other mechanical energy storage device  1124  provides a mechanical bias which tends to urge the deployable payload object  14   ep  away from contact with the vehicle body  14 . 
       FIG. 12  illustrates a notional deployment of a deployable payload object. In  FIG. 12 , the drag wings  610   a ,  610   b  of the deployable payload object  14   ep  are initially retained in place by a shape memory ring  1212 . Relaxation of the shape memory ring allows the drag wings to deploy and they spring open, whereupon the spin rate slows. When the drag wings are fully deployed, the spin energy decreases rapidly. Eventually, the deployed payload object  14   ep  alights on an alighting surface  1214 ; the drag wings are knocked off by the impact or by the shearing force of adjacent objects. The contact adhesive is exposed by the removed wings, and tends to cause the vehicle to stay at the site at which it alights. 
     The altitude of the vehicle above the local terrain can be established by the inclusion among the sensors in region  46  of  FIG. 1  of a downward-looking “radar” altimeter. Since the wavelengths of radar signals tend to be relatively large by comparison with the size of the vehicle, it is anticipated that a light-operated “lidar” equivalent will be used. The principles of lidars are well known in the art. The measured altitude can be transmitted to the remote controller or operator who can use the information to aid in control if the vehicle cannot be directly observed. As an alternative, the measured altitude can be provided to an autonomous navigation controller. The autonomous controller can use the altitude information from the lidar to increase or decrease the lift of the rotor disk in order to maintain a given altitude or to follow an altitude program. 
     While a downward-looking lidar altitude determination can give altitude, a side-looking lidar rangefinder can provide a rotating scan for determining proximity of surrounding objects such as trees or large rocks in an outdoor setting, or walls in an interior context. This simply requires correlating the scanned proximity information with the rotation rate which is known from the scanning optical camera. As an alternative, the scanning proximity signal can itself be used to aid in determining the rotation rate in much the same fashion as that for the scanning camera, by noting the recurrence rate of objects at particular distances. 
     One of the problems with existing reconnaissance flying vehicles is that of protecting them in the interim between initiation of a ground force action and the later time at which the flying vehicle is used during the operation. It will be appreciated that a flying vehicle must be protected against mud, water, and extreme shock notwithstanding that the particular trooper carrying the vehicle is exposed to these conditions. According to an aspect of the invention, the flying vehicle(s) is(are) protected by a common blister pack which is preferably hermetic or at least very tight. The blister pack may be individual, or may be a multiple pack such as those potable packs in which medical tablets or capsules are stored for later use.  FIG. 7A  shows the elements of a blister pack before closing, and  FIG. 7B  is a side elevation view of the structure of  FIG. 7A  in its assembled state. As illustrated in  FIG. 7A , a first piece  710  in the form of a sheet defines a set  710  of a plurality of open cavities or depressions  710   ca ,  710   cb ,  710   cc ,  710   cd , and  710   ce . A generally flat cover sheet  712  is dimensioned to cover the cavity(ies) of set  710 . The individual flyers  10   a ,  10   b ,  10   c ,  10   d , and  10   e  are placed in cavities  710   ca ,  710   cb ,  710   cc ,  710   cd , and  710   ce , respectively, and the cover sheet  712  is affixed to the flat portions of piece  710  with some kind of adhesive, illustrated as a surface  714 . As an alternative to the use of adhesive, the two pieces  710 ,  712  can be welded as by ultrasonic or heat, to for closed cavities. If desired, a bit of foam or other resilient material can be placed in each cavity before closing to restrain the flying vehicle against movement within the cavity. 
     While the thrust provided by nozzle or jet  270  of  FIG. 1A  is near the distal end or tip end  12   2  of the airfoil or wing  12 , those skilled in the art know that the nozzle or jet may be at a location lying between the tip end  12   2  and the root end  12   1 , but that less torque will be developed for a given thrust. Some embodiments of the vehicle  10  may include photovoltaic “solar” panels or cells for augmenting the on-board batteries with locally produced power. It might even be possible to recharge the batteries of a vehicle after a period of rest while exposed to sunlight. Another version might incorporate high-efficiency thermophotonic or thermophotovoltaic cells, positioned adjacent the fuel grains to take advantage of the heat generated therein during operation. 
     An apparatus for flight ( 10 ) according to an aspect of the invention comprises an airfoil ( 12 ) with an attached payload ( 14 ,  14   ep ), and propulsion means ( 210 ) associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the attached payload ( 14 ,  14   ep ), for thereby defining a rotor disk ( 99 ). The apparatus ( 10 ) also comprises physical means ( 510 ) for adjusting the lift of the airfoil ( 12 ), and control means ( 410 ,  510 ) coupled to the physical means ( 510 ) for causing the lift adjustment of the airfoil ( 12 ) to tilt the rotor disk ( 99 ). In one embodiment, the rotating airfoil ( 12 ) defines a leading edge ( 12   le ) and a trailing or lagging ( 12   te ) edge, and the physical means ( 510 ) comprises means ( 512 ,  514 ,  521 ,  522 ,  531   a ,  531   b , . . . ) for ejecting gas at a location ( 533 ) near the trailing edge ( 12   te ) in at least one plane ( 598   u ,  598   d ) which does not coincide with, or is skewed relative to, the plane of the rotor disk ( 99 ). In a preferred embodiment, the means ( 512 ,  514 ,  521 ,  522 ,  531   a ,  531   b , . . . ) for ejecting gas includes means for ejecting gas in a generally periodic manner in the plane not coincident ( 598   u ,  598   d ) with the rotor ( 99 ). In one embodiment, the airfoil ( 12 ) of the apparatus is elongated, and defines a distal end ( 12   2 ) remote from the payload ( 14 ,  14   ep ), and the propulsion means ( 210 ) comprises a solid-fuel powered bypass jet, with an exhaust directed ( 12   t ) generally perpendicular to an axis ( 8 ) of the elongation of the airfoil ( 12 ). In one version, the exhaust is directed generally in the plane of the rotor disk ( 99 ). The transverse location of the exhaust may lie generally between the distal end ( 12   2 ) of the airfoil ( 12 ) and the payload ( 14 ,  14   ep ), or it may be at the distal end ( 12   2 ) of the airfoil ( 12 ). 
     According to another aspect of the invention, an apparatus ( 10 ) for moving a load in a selected direction comprises an airfoil ( 12 ) with a fixedly attached payload ( 14 ,  14   ep ), and propulsion means ( 210 ) associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the attached load ( 14 ,  14   p ), for thereby rotating the airfoil ( 12 ) to define a rotor disk ( 99 ). Physical means ( 510 ) are provided for adjusting the lift of the airfoil ( 12 ). Control means ( 410 ,  510 ) are coupled to the physical means ( 510 ) for causing the lift adjustment of the airfoil ( 12 ) to tilt the rotor disk ( 99 ) in a manner which moves the airfoil ( 12 ) with the attached load in the selected direction. 
     A flying apparatus ( 10 ) according to another aspect of the invention is for moving a load ( 14 ,  14   p ). The apparatus ( 10 ) comprises an airfoil ( 12 ) with an attached load ( 14 ,  14   p ) fixed to the airfoil ( 12 ), and propulsion means ( 210 ) associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the attached load ( 14 , 14   p ) together, for thereby defining a rotor disk ( 99 ). Physical means ( 510 ) are provided for adjusting the lift of the airfoil ( 12 ), and control means ( 410 ,  510 ) are coupled for causing the lift adjustment of the airfoil ( 12 ) to provide at least one of collective and cyclic control. 
     A flying apparatus ( 10 ) according to a further aspect of the invention comprises an airfoil ( 12 ) with an attached load ( 14 , 14   p ) adjacent a first end of the airfoil ( 12 ), and a jet ( 270 ) lying between the first ( 12   1 ) and second ( 12   2 ) ends of the airfoil ( 12 ) for rotating the airfoil ( 12 ) and attached load ( 14 ,  14   ep ). 
     An apparatus ( 10 ) for flight according to an aspect of the invention comprises an airfoil ( 12 ) with a payload ( 14 ,  14   ep ) which is fixed to the airfoil ( 12 ), and propulsion means ( 210 ) associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the payload ( 14 ,  14   ep ), thereby defining a rotor disk ( 99 ). The airfoil ( 12 ) with payload ( 14 ,  14   ep ) fixed thereto has no attached payload ( 14 ,  14   ep ) which rotates at a rate other than the rotation rate of the airfoil ( 12 ). Physical means ( 510 ) adjust the lift of the airfoil ( 12 ), and control means ( 410 ,  510 ) are coupled to the physical means for causing the lift adjustment of the airfoil ( 12 ) to tilt the rotor disk ( 99 ). 
     An apparatus ( 10 ) for flight comprises an airfoil ( 12 ) with a fixedly attached body ( 14 ), where the airfoil ( 12 ) and fixedly attached body ( 14 ) together defining a center of mass. The attached body includes payload attachment means ( 810 ,  812 ;  910 ,  1016 ,  1120 ) for attaching a payload ( 14   ep ) centered on the center of mass ( 8 ), which payload ( 14   ep ), when attached, is fixedly attached to the body ( 14 ). A payload ( 14   ep ) is coupled to the payload attachment means ( 810 ,  812 ;  910 ,  1016 ,  1120 ), and propulsion means ( 210 ) are associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the fixedly attached body ( 14 ), for thereby defining a rotor disk ( 99 ). Physical means ( 510 ) are provided for adjusting the lift of the airfoil ( 12 ). Control means ( 410 ,  510 ) are coupled to the physical means for causing the lift adjustment of the airfoil ( 12 ) to tilt the rotor disk ( 99 ), and control means are provided for controlling the payload attachment means ( 810 ,  812 ;  910 ,  1016 ,  1120 ) for disengaging the body ( 14 ) from the payload ( 14   ep ) at selected one of (a) time and (b) location. 
     An apparatus ( 10 ) for flight according to an aspect of the invention comprises an elongated airfoil ( 12 ) with an attached payload ( 14 ,  14   ep ), which airfoil ( 12 ) defines a longitudinal axis ( 8 ). Propulsion means ( 210 ) are associated with the airfoil ( 12 ) for rotating the airfoil ( 12 ) and the attached payload ( 14 ,  14   ep ), for thereby defining a rotor disk ( 99 ). The propulsion means ( 210 ) comprises means ( 210 ,  224 ,  230 ,  252 ,  266 ) for generating gas under pressure and means ( 270 ) for releasing the gas under pressure in a direction ( 12   t ) generally tangent to a radius of the rotor disk ( 99 ) and from a location near an end ( 12   2 ) of the airfoil ( 12 ). The propulsion means ( 210 ) in one embodiment of this aspect of the invention comprises an ejector ( 230 ,  266 ) driven by a fuel grain ( 310   a ,  310   b , . . . ), and the fuel grain ( 310   a ,  310   b , . . . ) may generate hot gas which is partially combustible. In another embodiment according to this aspect of the invention, the means ( 210 ,  224 ,  230 ,  252 ,  266 ) for generating gas under pressure comprises a fuel grain ( 310   a ,  310   b , . . . ) which, in operation, creates partially combustible hot gas under pressure, and a first ejector ( 230 ) into which the partially combustible hot gas under pressure is introduced, for mixing the partially combustible hot gas with atmospheric oxygen ( 228 ), to generate hot combusted gas. A second ejector ( 266 ) receives the hot combusted gas, and heats atmospheric gas ( 264 ) to generate the gas under pressure (GUP). 
     A protective package ( 710 ,  712 ) according to another aspect of the invention is for individually protecting flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ). The protective package ( 710 ,  712 ) comprises a first piece ( 710 ) defining a cavity (set  711 ) in which each cavity is larger in length, width and depth than corresponding dimensions of the flying vehicle ( 10 ). A second piece ( 712 ) is provided having dimensions sufficient to occlude or close off the entirety of the cavity(ies) (set  711 ). Means ( 714 ) are provided for affixing the second piece ( 712 ) to the first piece ( 710 ) so as to define with the cavity (set  711 ) a closed package containing the flying vehicle. The means may be adhesive or welding, rivets, or any other attachment. In one embodiment, at least one of the first plastic piece and the second piece ( 712 ) is transparent. Either the first piece ( 710 ) or the second piece ( 712 ), or both, may be of a plastic material. The second piece ( 712 ) may be monolithically hinged to the first plastic piece to define a clamshell. 
     A protective package ( 710 ,  712 ) according to another aspect of the invention is for accommodating a plurality of flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ) and includes a first generally planar piece ( 710 ) defining a plurality, equal in number (five in the example of  FIGS. 7A and 7B ) to the number of the plurality of flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ), of individual open cavities ( 710   ca ,  710   cb ,  710   cc ,  710   cd ,  710   ce ). Each of the open cavities ( 710   ca ,  710   cb ,  710   cc ,  710   cd ,  710   ce ) is dimensioned to accommodate one of the flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ). The package also includes a second piece ( 712 ) dimensioned to occlude the plurality of individual open cavities. The second piece ( 712 ) is applied to the first piece ( 710 ) to occlude the open cavities and thereby define the plurality of closed cavities. Each of the closed cavities accommodates one of said flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ). 
     A method for storing flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ) according to a further aspect of the invention comprises the step of encapsulating each flying vehicle in shrink-wrap film, and heating the shrink-wrap film to cause the film to shrink about the flying vehicle. 
     Another method for storing flying vehicles ( 10   a ,  10   b ,  10   c ,  10   d ,  10   e ) according to an aspect of the invention comprises the steps of placing a flying vehicle in each cavity (set  711 ) of a sheet ( 710 ) defining plural cavities (set  711 ), and applying a single second sheet ( 712 ) over the open side of the plural cavities (set  711 ) to form sealed chambers, each holding one flying vehicle.