Abstract:
An apparatus for defending a vehicle against an approaching threat is disclosed. The apparatus comprises a body deployable from the vehicle having at least one reference axis thereon. A plurality of spaced divert thruster means are associated with the body for translating the body relative to the threat upon deployment from the vehicle. A plurality of spaced attitude control thruster means are associated with the body for providing roll, pitch and yaw to orient the reference axis to a desired orientation relative to the threat. A guidance system, associated with the reference axis, is provided for communicating electronic signals to the attitude control thruster means and divert thruster means for controlling the position of the body relative to the threat to eliminate the threat. The guidance system preferably operates independently of the vehicle after deployment.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to defense mechanisms for atmospheric and extraterrestrial vehicles. More particularly, the present invention relates to an apparatus releasable from the vehicle which can either serve as a decoy, by diverting an approaching threat, and/or serve as a kinetic kill device by positioning itself to collide with the threat. 
     2. Description of the Related Art 
     Defensive strategies used by aircraft to defend against approaching threats include passive defense to avoid a threat, active defense to deceive a threat, and lethal defense to kill the source of the threat. Methods currently being used to implement these defensive strategies include low altitude sanctuary, terrain masking, expendable countermeasures, electronic countermeasures (ECM), air-to-air missiles, air-to-surface missiles, and most recently, low observable, or “stealth” technology. These methods provide a substantial measure of survivability in a very hostile threat environment, but face continual pressure to keep pace with threat upgrades. Examples of prior art decoys/targets are included in the following references: 
     U.S. Pat. No. 2,957,417, issued to D. D. Musgrave, entitled “Missile Decoy”, discloses a decoy system comprising a shroud fixed to the fins of a missile having an internal reflective surface of partial parabolic shape. 
     U.S. Pat. No. 2,898,588, issued to C. L. Graham, entitled “Attack Deviation Device”, discloses a device comprising a hollow elongated member which is towed behind an aircraft. There is structure in the member for reflecting radar signals impinging thereon. 
     U.S. Pat. No. 4,808,999, issued to D. Toman, entitled “Towed Decoy with Fiber Optic Link”, discloses a decoy which is towed behind an aircraft using a tow line which incorporates a fiber optic link through which signals are transmitted. 
     U.S. Pat. No. 4,419,669, issued to D. M. Slager, entitled “Controlled Scintillation Rate Decoy”, discloses a decoy having a spherical body with microwave reflectors for reflecting incident radar energy in a manner to provide the decoy with the selected radar cross section. 
     U.S. Pat. No. 3,290,681, issued to R. H. Beteille, entitled “Device for Jamming Radar Detection and Interception of Ballistic Missiles”, discloses the use of decoys consisting of metallized inflatable balloons. 
     U.S. Pat. No. 3,568,191, issued to J. C. Hiester, entitled “Method for Defending an Aircraft Against a Frontal Attack”, discloses a method of launching a rocket in the direction of travel of an aircraft, the rocket containing, in a collapsed, condition a collapsible reflector having three orthogonal surfaces of wire mesh. The reflector is ejected from the rocket, expanded, and thereby towed, the reflector serving as a target for attracting the interceptor. 
     U.S. Pat. No. 3,126,544, issued to W. H. Greatbatch, Jr., entitled “Method of Deception for an Aircraft”, discloses deployment of a spherically shaped decoy that is provided with a plurality of target plates. The decoy is illuminated with RF energy, the interceptor being deceived into following the decoy target. 
     U.S. Pat. No. 3,045,596 entitled “Guided Missile”, issued to R. S. Rae discloses a controlled missile having structure including a spherical body having spaced ports for the discharge of jets in such a manner so as to lift and propel the missile. The guided missile is launched and controlled in aerial flight by the operator from the time of launching to the time of contact with a target. The Rae device weighs on the order of 300 pounds and is two feet in diameter. It carries explosives. 
     None of the aforementioned references disclose a defense system or apparatus having 1) a guidance system capable of operating independently of the launch vehicle and/or operator, and 2) a propulsion system having the ability to accurately maneuver the body relative to the threat in the absence of a substantial axial velocity so as to eliminate the threat. 
     As will be disclosed below, present applicants have developed a novel apparatus for defending a vehicle against an approaching threat which provides these capabilities. This patent application is related to patent application Ser. No. 07/491,798, entitled “Killer Volleyball Defense System” filed concurrently herewith and patent application Ser. No. 07/493,087, entitled “Killer Volleyball Launcher”. All three patent applications are assigned to the present assignee, Rockwell International Corporation. 
     The term “Killer Volleyball” or “KV”, as used herein and in the aforementioned, concurrently filed patent applications, refers to the presently disclosed and claimed apparatus for defending a vehicle. This is the terminology used during its development by Rockwell International Corporation. 
     OBJECTS AND SUMMARY OF THE INVENTION 
     It is a principal object of the present invention to defend a vehicle from an approaching threat without the need for a propulsion system which provides a substantial axial velocity in the direction of the threat. 
     Another object is to defend the vehicle from a threat which may be approaching from the aft sector. 
     Another object is to defend an atmospheric vehicle from an approaching airborne threat. 
     These and other objects are achieved by the present invention which is an apparatus for defending a vehicle against an approaching threat. In its broadest aspects, the apparatus (i.e. Killer Volleyball) comprises a body deployable from the vehicle having at least one reference axis thereon. A plurality of spaced divert thruster means are associated with the body for translating the body relative to the threat upon deployment from the vehicle. A plurality of spaced attitude control thruster means are associated with the body for providing roll, pitch and yaw to orient the reference axis to a desired orientation relative to the threat. A guidance system, associated with the reference axis, is provided for communicating electronic signals to the attitude control thruster means and divert thruster means for controlling the position of the body relative to the threat to eliminate the threat. The guidance system preferably operates independently of the vehicle after deployment. 
     In its narrower aspects, the body is substantially spherical with a uniformly textured outer surface. It is particularly adaptable for use with atmospheric vehicles being defended against airborne threats. In one embodiment the guidance system includes a seeker means for tracking the threat, the seeker means providing electrical information for the guidance system to maneuver the apparatus for a collision with the threat, thereby providing a kinetic kill and elimination of the threat. In another embodiment the apparatus serves as a decoy, emitting a signal for diverting the threat away from the vehicle, thereby eliminating the threat. 
     Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 illustrates the overall engagement scenario for a KV aircraft defense. 
     FIG. 2 is a perspective view of the KV apparatus. 
     FIG. 3 is a perspective view, partially broken away, of an aft end of an aircraft having a KV launcher and a plurality of KVs mounted therein. 
     FIG. 4 is a functional schematic diagram of a threat warning system and KV guidance subsystem. 
     FIG. 5A is a cutaway plan view of the KV along the plane of the seeker axis to show propulsion component and avionics layout. 
     FIG. 5B is a view of the KV taken along line  5 B— 5 B of FIG.  5 A. 
     FIG. 5C is a view of the aft end of the KV illustrating the primary structure and externally mounted propulsion components. 
     FIG. 5D is a view taken along line  5 D— 5 D of FIG.  5 C. 
     FIG. 5E is a view taken along line  5 E— 5 E of FIG.  5 C. 
     FIG. 6 is a schematic illustration of the KV propulsion system. 
     FIG. 7 is a schematic illustration of the KV environment, illustrating aerodynamic considerations. 
     FIG. 8 illustrates the aerodynamic symmetry of a spherical KV. 
     FIG. 9 is an enlarged perspective view of a preferred launcher apparatus for multiple deployment of KVs. 
     FIG. 10 is an exploded perspective view of the launcher illustrated in FIG.  9 . 
     FIG. 11 illustrates a decoy engagement scenario for a decoy embodiment of the KV. 
    
    
     The same elements or parts throughout the figures of the drawings are designated by the same reference characters, while equivalent elements bear a prime designation. 
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     The overall engagement scenario for Killer Volleyball (KV) aircraft defense includes three general phases: threat detection, KV launch, and threat intercept. 
     Referring to the drawings and the characters of reference marked thereon, FIG. 1 illustrates this engagement scenario. 
     In threat detection, the target aircraft  10  recognizes an approaching threat  12  by means of a threat warning system (TWS). At this time in the scenario the KV  14  is supported by the aircraft. As will be more fully described below, the TWS is capable of detecting the threat  12  at a sufficient distance to permit adequate reaction and provides the data necessary to determine the proper timing for KV  14  release. The TWS data is also used to provide situational awareness (SA) to the aircrew and possible participation in the KV  14  employment decision. 
     Once the threat  12  has been detected and the proper timing for KV  14  release determined by the fire control system, the KV  14  will be launched from the aircraft  10 . Target lock-on against the attacking threat  12  will then occur at a safe distance behind the aircraft, releasing the KV  14 . The launch is timed to permit a threat intercept within the KV operational envelope  16  and at a safe distance from the aircraft  10  being defended. 
     Referring now to FIG. 2, a perspective view of the KV  14  is illustrated. The KV  14  includes, inter alia, divert thrusters  18 , attitude control thrusters  20 , and a seeker  22 . The divert thrusters  18  are located in a divert plane, z-y, normal to the seeker viewing axis x (i.e. reference axis). 
     To intercept the threat, the KV  14  uses its divert thrusters to position itself along a trajectory  24  (see numeral designations  14 ′,  14 ″ in FIG. 1) for a collision with the threat  12 . This guidance is independent of the host aircraft  12  and is based on data provided by the KV seeker  22 . Destruction of the threat (represented by numeral designation  26 , in FIG. 1) is achieved by a kinetic collision between the KV  14  and the threat  12 . The KV  14 , in its preferred embodiment, does not require explosives or fuses. 
     The preferred technique used to intercept the threat involves two major elements: deceleration of the KV  14  parallel to the longitudinal axis of the aircraft  10 , and translation in a plane z-y, perpendicular to the seeker axis x. 
     Prior to launch the KV  14  is secured within a launcher, designated generally as  28 , in FIG.  3 . The launcher  28  is preferably located at an aft portion of the aircraft  10 , each seeker viewing axis, x, being parallel to the longitudinal axis of the aircraft x. 
     Following launch from the aircraft  10 , the KV  14  decelerates rapidly due to aerodynamic drag. This deceleration provides a rapid separation between the aircraft  10  and KV  14 . It also provides rapid closure with the threat  12  and reduces the KV  14  flight time required for intercept. While the KV  14  is decelerating longitudinally, its divert thrusters  18  are used to translate the KV  14  laterally and vertically in the divert plane, z-y, perpendicular to the seeker axis x. This plane of translation, z-y, defines a barrier to the approaching threat  12 . 
     Referring now to FIG. 4, the threat detection, tracking and guidance functions and their interrelationships are illustrated in terms of the host platform, or launcher  28 , and the KV  14 . For threat detection, a search and track system  30  is provided for detecting an approaching threat and analyzing information regarding the threat&#39;s angular position. The analyzed information is input to a range measurement unit  32  for determining relevant distances and closure rates. Search and track system  30  and range measurement unit  32  cooperate to reduce false alarms. 
     The search and track system  30  may be, for example, a passive IR search and track set or an active radar set. The search and track system  30  must provide coverage of a solid angle appropriate to the anticipated threat volume and should be capable of handling multiple targets. The range measurement unit  32  may be a laser range finder or low probability of intercept (LPI) radar. These permit location of the threat in three dimensions and accurate determination of the trajectory and/or closing rate for use in determining an optimum launch time for the KV  14 . If an active sensor is not provided, approximate ranges and closing rates of incoming threats can be determined utilizing passive ranging techniques such as measuring the rate of change of the IR source intensity and of angular position, to estimate the threat trajectory. 
     The range measurement unit  32  provides range and closure rate information to the launch computer  34 . (The launch computer  34  and range measurement unit  32  form the fire control system.) The launch computer  34  determines the optimum time for a launch and transfers the appropriate signal to the launcher  36 . The launch computer  34  also provides information on system status to the host avionics  38 . 
     The angular position information is provided by the search and track system  30  to the KV seeker  22 . Target assignment and, preferably, range information is passed from the launch computer  34  to the KV guidance computer  34  prior to launch. Boresight correlation may be provided between the seeker  22  on the KV  14  and the search and track system  30  to handle multiple targeting assignments. 
     Information relating to the angular offset of the threat from the normal to the KV thrust (divert) z-y plane is provided by the seeker  22  to the guidance computer  40 . The Inertial Measurement Unit (IMU)  42  provides 3-axis attitude information to the seeker  22  and guidance computer  40 . The guidance computer  40  determines the optimum trajectory for intercept and provides guidance information to the thrust control  44  which includes the attitude control thrusters  20  and divert thrusters  18 . 
     The functional arrangement of the components in FIG. 4 may be modified to suit the particular capabilities of the threat warning system which is installed on board a particular host platform. For example, the threat warning system may not include a fully capable search and track system. However, it may include a tail warning sensor capable of providing coarse angular position or sector, with or without range information, that can be transferred to the KV seeker and guidance computer, as appropriate. In this case the KV seeker must have a relatively wide field of view at least equal to the angular uncertainty of the information provided by the tail warning sensor. 
     If range information is provided, a launch computer may be included in the host platform threat warning system for computing the best time to launch, as in the system of FIG.  4 . Alternatively, this computer may be omitted and the computation provided by the guidance computer on board the KV. In other respects the operation of the threat warning and KV guidance systems is, as described in connection with FIG.  4 . 
     In the event the host platform has no threat warning system and cannot accommodate the addition of such a system, the threat detection and tracking function may be accomplished on board the KV. In this case the sensor on board the KV must perform the search and track functions as well as the seeker guidance function. This requires a very wide field-of-view sensor with high resolution, approaching as nearly as possible (within size and monetary constraints) the performance of the search and track system included in the system of FIG.  4 . 
     As an alternative to the above system configuration, where the host platform has no threat warning system, and/or the KV launcher is not accommodable by the host platform, the complete system represented in FIG. 4 or any modification of same may be installed in a pod to be mounted on the host platform. 
     Referring now to FIGS. 5A,  5 B and  5 C, a preferred embodiment of the KV apparatus, designated generally as  45 , is illustrated. The body  45  includes a two-piece spherical shell  46  preferably formed of aluminum or a composite material. The shell  46  has a uniformly textured outer surface having roughly equidistantly spaced dimples. The avionics and propulsion components are packaged by mounting the same to an internal cylindrical primary structure  48 . 
     The ends of the primary structure  48  are connected to the shell  46  by threaded fasteners  50  (threaded or riveted). The two shell halves may be fastened, for example, using rivets or threaded fasteners. 
     The avionics include a seeker  52  (also illustrated as seeker  22  in FIG.  2 ), an inertial measurement unit (IMU)  54  ( 42  in FIG.  4 ), a processor  56 , a battery  58 , a power distribution device  60 , a valve driver  62 , and an umbilical disconnect  64 . There is no telemetry on board, since, in the preferred embodiment the KV operates autonomously once released from the launcher. The seeker  52  is mounted to the primary structure  48 , with the seeker viewing axis x pointing aft relative to the ballistic path of the jettisoned KV. The IMU  54  is mounted in close proximity to the seeker  52  to minimize vehicle attitude error due to structural deflection. The umbilical disconnect  64  is mounted at the opposite end of the primary structure  48  from the seeker  52 , such that the ejection of the KV  45  from the host aircraft disengages the umbilical  64  while pointing the seeker  52  in the aft direction, x. The other avionic components are mounted to the primary structure  48  to balance the vehicle such that its center-of-mass lies at the centerpoint of the spherical shell, i.e. the “KV centerpoint”. 
     As best illustrated in FIG. 5B, four divert thrusters  66  (also illustrated as divert thrusters  18  in FIG. 2) located in a divert plane z-y normal to the seeker viewing axis x are mounted on the outside periphery of the primary structure  48  pointing radially outward at each 90° location. Because the center-of-mass of the KV  45  coincides with the KV centerpoint, the thrust vector of the divert thrusters  66  act through the center-of-mass, minimizing disturbing torques to the KV due to thrust vector/center-of-mass offset. 
     Additionally, the aerodynamic center of the KV coincides with the KV centerpoint, regardless of the direction of its flight. This minimizes disturbance torques due to aerodynamic load vector/center-of-mass offset. 
     The dynamic masses, i.e. the propellants and pressurant, are also arranged such that their center-of-masses coincide with the KV centerpoint during utilization of the propellants. The four spherical propellant tanks  68 ,  70  are equally spaced between the divert thrusters  66  on the outside periphery of the primary structure  48 . The two fuel tanks  68  and the two oxidizer tanks  70  are mounted opposite each other, respectively, to offset the dynamic center-of-mass shift due to propellant usage. 
     The center of the spherical pressurant tank  72  coincides with the KV centerpoint, which eliminates center-of-mass shift due to usage of the pressurant, since the center-of-mass of the pressurant is distributed uniformly about the center-of-mass of the KV throughout the mission. The pressurant is stored in the pressurant tank  72 , and after initiation progressively flows to the four propellant tanks  68 ,  70 , which are symmetrically mounted about the center-of-mass. 
     Pitch, yaw and roll stability is afforded by six attitude control system (ACS) thrusters  74 , which are mounted in a plane forward of the divert thrusters  66 . The horizontally-oriented ACS thrusters provide roll and yaw control, while the vertically-oriented thrusters provide pitch control. 
     The bipropellant, pressure-fed propulsion system is shown schematically in FIG.  6 . The propellants are nitrogen tetroxide (NTO) as the oxidizer and monomethyl hydrazine (MMH) as the fuel. This provides hypergolic ignition. Helium is the pressurant. 
     The pressurant tank  72  is loaded to 10,000 psi through the service valve  76 . The pressurant is contained entirely within the pressurant tank  72  until the pyrotechnically actuated helium isolation valve  78  is opened. Actuating the helium isolation valve  78  shears a metallic nipple to allow helium to flow to the regulator  80 . The regulator  80  provides a constant pressure to the helium manifold  82 , propellant tanks  68 ,  70  and divert thrusters  66 . 
     Each propellant tank  68 ,  70  has a metallic diaphragm, which contains the propellant within the tank and prevents the propellants from mixing in the helium manifold  82 . The helium pressure acting on the diaphragm forces the propellant through the tank mount  84  and into the propellant manifolds  86  and  88 . As the propellant is utilized through the divert thrusters  66  and ACS thrusters  74 , the diaphragms reverse to positively push the propellant out of the tanks. 
     The tank mount  84  provides a means to bolt the propellant tank to the primary structure  48 . Each tank mount  84  houses a service valve  90 , isolation valve  92 , and filter. The service valve  90  is provided for loading propellant into the propellant tank  68 ,  70 . The isolation valve  92  contains the propellant in the propellant tank  68 ,  70  until pyrotechnically actuated, after which the propellant is free to flow into the propellant manifolds  86  and  88 . The filter prevents particulate contamination from entering the propellant manifolds  86 ,  88 . 
     The helium manifold  82 , oxidizer manifold  86  and fuel manifold  88  are integral channels imbedded circumferentially within the wall of the primary structure  48 . Passages into and out of the manifolds are provided by ports through the outer surfaces of the structure into the manifold. The divert thrusters, tube flanges, and propellant tanks are bolted onto the flat surfaces on the periphery of the primary structure  48 , with O-rings to provide sealing between the components and the structure. 
     The divert thrusters  66  and ACS thrusters  74  have trim orifices  94 ,  96 , respectively, to calibrate the flow rate of propellants into the thrusters. The trim orifices  94  for the divert thrusters  66  are located in the mounting face of the divert thrusters  66  and are sealed with o-rings. The trim orifices  96  for the ACS thrusters  74  are welded into the inlet lines of the ACS thrusters  74 . 
     The thrusters are fired on demand for divert maneuvers and attitude control, and are capable of steady-state or pulse-mode operation. The thrusters are fired by electrically energizing the solenoids of the individual thruster valves. These valves are spring-loaded closed while the solenoids are in the unenergized state. 
     The bipropellant thrusters are capable of on-off operation at rates as fast as 100 hertz and have a rapid thrust rise time of less than 5 milliseconds to 90 percent rated thrust level. Maximum thrust of the described divert thrusters is preferably 170 lbs. 
     Service valves  98 ,  100  and  102  are utilized for leak checking the helium, oxidizer and fuel manifolds,  82 ,  86 ,  88  respectively. 
     The seeker  52  provides for transmission of the optical data to its focal plane. The focal plane data is sent to processor  56 , along with the vehicle attitude data from the IMU  54 , to determine the vehicle flight corrections and maneuvers required to intercept the threat. The corrections are relayed to the valve driver  62 , which commands the appropriate divert and ACS thruster valve(s)  66 ,  74  open. The data is reevaluated and new commands are given, nominally, every 10 milliseconds. 
     The battery  58  provides the on-board power for the electric and electronic devices. The power distribution device  60  provides properly conditioned power for each electrical component&#39;s requirements. 
     The spherical shell  46  preferably has a diameter in a range of between ten and fourteen inches. In an embodiment with a twelve inch diameter, the attitude control thrusters may be mounted in a plane approximately 4.7 inches forward the divert plane to provide a moment arm for pitch and yaw control. The approximate weight range of the KV is between 15 and 25 pounds. 
     Referring now to FIG. 7, a schematic illustration of the KV environment illustrating aerodynamic considerations is shown. 
     The kinematics of the trajectory to successfully achieve a collision with the threat and the basic problems of launching any device from an aircraft  106  introduces a kinetic kill device to a violently turbulent environment comprised of the aircraft wake  108 , jet exhaust  110 , and tip vortices  112  as well as a wide range of potential wind incidence angles. These effects produce potentially large forces and moments on the KV  14 , which could exhaust a limited fuel supply. 
     The turbulence may be divided into two characteristic components, a uniform gust intensity field and a rotational or velocity shear flow. The uniform gust contributes to the sizing of the divert thrusters, and the rotational gust sizes the attitude stabilization thrusters used to maintain the angular orientation of the thrust plane and seeker axis relative to the earth reference system. 
     In order to minimize the size and weight of the device, the stabilization requirements due to these disturbances must be minimized. Referring to FIG. 8, it is illustrated that a spherical shape is preferred for the KV  14 , due to a sphere&#39;s inherent aerodynamic symmetry (see lines  116  of symmetry). The static moments generated by a uniform flow field are minimized with a sphere as compared to the natural tendency for an arbitrary shape to weather vane or tumble. Since the device is required to laterally and vertically translate to position for a kinetic intercept while decelerating in the longitudinal axis, a very large range of wind incidence angles are possible. 
     Due to the presence of the thruster nozzle orifices on the surface of the sphere, additional moments may be generated increasing the required thrust for stabilization. A uniformly rough surface texture minimizes this effect. 
     The size of the vehicle determines the aerodynamic drag and therefore the separation characteristics and the limits of its lateral envelope. As drag increases, the time aloft is reduced, but the lateral envelope is also reduced. Sizing of the divert thrusters and fuel requirements depend upon the size of the device and the end-game requirements for collision. The sizing of the attitude stabilization thrusters is determined by the wake disturbance and aerodynamic stability and control requirements. A uniformly rough surface offers the best solution for minimizing stability and control requirements which, in turn, reduce the size and weight of the device. 
     Referring back now to FIG.  3  and also to FIG. 9, a launching apparatus  28  is illustrated which is preferably attached to a structural support (not shown) within the host aircraft. The host aircraft  10  includes an external fairing  118  with an aft opening which provides a field of view for a portion of the launching apparatus  28 , as will be more fully described below. 
     The apparatus  28 , for launching, includes a rotary launcher  120  attached to the host aircraft  10 . The launcher  120  has a plurality of support fittings  122  thereon. Each support fitting  122  is used for mounting a respective launch tube assembly  124 . A plurality of projectile retention means  126  are provided for securely retaining each KV  128 , previously illustrated as KV  14 , in FIG. 1, within its respective launch tube assembly  124 . A plurality of ejection cartridges  130  are provided, each being mounted to a respective launch tube assembly  124 . Each ejection cartridge interfaces with a respective projectile retention means  126  to provide an ejection thereof. 
     The rotary launcher includes a central rotatable hub  132  mounted in the host aircraft. The support fittings  122  extend radially outward from the hub  132 . As can be seen in FIG. 10 the rotary launcher  120  also includes a drive motor  134  mounted within the central rotatable hub  132 . Each launch tube assembly  124  includes a rigid shell  136 , a shell mount  138 , a cartridge housing  140 , and an umbilical connector  142 . The rigid shell  136  contains the retention means  126  and KV  128 . Each shell  136  has stabilization races  144  formed on an inner surface thereof. The shell mounts  138  support each rigid shell  136  to its respective support fitting  122 . Each cartridge housing  140  is mounted on a rear portion of a shell  136  for containing the ejection cartridge  130 . The umbilical connectors  142  provide KV/host aircraft electrical and cooling interfaces. 
     Each projectile retention means  126  is preferably formed of a rigid foam having a thermal protective coating. The foam partially encapsulates the KV and secures it to the launch tube assembly  124  while distributing the thrust load from the ejection cartridge  130 . The rigid foam provides a pressure seal that contains ejection cartridge gas until an adequate pressure is obtained to propel the KV  128  and retention means  126  at the prescribed velocity. The rigid foam may be, for example, plastic, teflon, or ceramic. The coating may be, for example, metal foil or a ceramic. The foam is preferably in sections  146  which separate forming a sabot when the KV is launched. Each foam section  146  includes at least one stabilization track  148  for mating with an associated stabilization race  144  on the shell. Each ejection cartridge  130  is preferably an interchangable electrically ignited pyrotechnic device or an electrically actuated compressed gas cartridge, providing a large volume of gas to propel the projectile  128  at the prescribed velocity. 
     As noted, the host aircraft  10  includes an external fairing  118  with an aft opening. During operation, the rotary launcher  28  rotates to position a launch tube assembly  124  in a position adjacent to the aft opening to allow sequential projectile launch. Each launch tube assembly  124  is of sufficiently short length to provide a 60 degree conical field-of-view relative to a center of the KV. The KV is preferably deployed in a manner parallel to the aircraft line of flight. 
     Obviously, other modifications and variations of the present invention are possible in light of the above teachings. It is, therefore, to be understood that, within the scope of the appended claims, the invention may be practiced otherwise than as specifically described. 
     For example, although the present inventive concepts have been described in detail with respect to a seek and kill scenario, these principles may be readily extendable to application of the KV as a decoy, for example to counter SAMs and interceptors. Referring now to FIG. 11, in a decoy scenario the KV apparatus  150  is equipped with means for emitting a signal  152  for diverting the threat  154  away from the vehicle  156 , thereby eliminating the threat. In this instance, the requirement of a seeker is obviated. 
     In one decoy embodiment, the KV is outfitted with a broadband RF repeater. A threat  154  using radar  155  guidance is pulled off the protected aircraft  156  by a stronger RF return emanating from the KV  150 . In this instance, the KV is programmed to fly away from the host aircraft, as illustrated by numeral designations  150 ′,  150 ″ 0  eliminating the chance of the missile flying through the decoy into the protected aircraft. 
     In another embodiment, in order to protect the host aircraft from an infrared seeking threat, the KV is modified to produce a high IR signature while flying away from the host. The IR missile loses track of the host vehicle and closes in on the KV. 
     In yet another decoy embodiment, the signal is emitted by reflection of a signal received by the threat. 
     It is to be understood that in the seek and kill scenario, it is not necessary that a kill be provided solely by a kinetic collision. For example, the KV may be equipped with explosives or other disabling means such as nets for obstructing the progress of the threat. 
     Furthermore, it is understood that divert propulsion is not limited to that which is solely perpendicular to the reference axis. Additional thrust means may be added to provide the body an axial velocity in the direction of the threat, if necessary, to improve the efficiency of the device. However, it is understood that the threat velocity and not any KV axial velocity is the primary mechanism for providing closure between the KV and the threat. 
     With respect to the launching system mentioned, although a rotary launcher has been illustrated, other launchers such as linear launchers may be used. In this respect launch tube assemblies may be arranged in a linear fashion for installation on smaller host aircraft. An alternative to using the proposed ejection cartridge launch mechanism for KV launcher separation may include use of the KV divert thrusters or aerodynamic pressure to achieve such a separation. 
     U.S. bombers can benefit greatly from a KV defensive system. Installed on a launcher and integrated with the existing aircraft defense systems, the KV can significantly improve bomber survivability. 
     A key requirement for tactical aircraft (both fighters and ground attack aircraft) is to be able to operate and release weapons at medium altitudes because of increased targeting opportunity and manufacturing flexibility in this altitude region. However, this is the flight regime with the highest potential of attacks from interceptors and SAMs. The defensive capability of the KV system would allow tactical aircraft (which otherwise could not perform combat operations at medium altitude because of unacceptable attrition) to achieve less restricted medium-altitude air combat capability. The resulting improvement in mission effectiveness of tactical air forces would be significant. 
     SOF aircraft such as the AC-130 U are typically slow and unable to maneuver against a threat. They are also large and difficult to mask against sophisticated electronic detection and tracking technologies available globally. The KV system allows these aircraft to defend themselves without engaging overt hostilities while over flying politically sensitive areas, thus, improving the survivability and operational utility of these air vehicles. 
     The benefits to cargo aircraft are similar to SOF aircraft. Inter-theater airlifters such as the C-141 and C-5A would benefit from KV defensive systems during a high intensity conflict. A KV defensive system for intra-theater and medium haul airlifters that deliver cargo and troops into “hot” areas would be beneficial for the enhanced survivability and extending the useful penetration range of these aircraft. 
     A KV defensive support system would be useful for installation in executive aircraft, such as Air Force 1, protecting key personnel from a terrorist situation. 
     Although application of the KV concept has been described with respect to atmospheric vehicles, it may also be used for extraterrestrial vehicles. For example, it may be used to protect satellites, space stations, space shuttles and other extraterrestrial vehicles. 
     Furthermore, it has naval applications and application to ground forces. For example, the U.S. Navy has off-board IR and RF jammers. For those threat systems not defeated by these passive system, some form of active lethal defense is required. The killer volleyball concept presents a solution to this problem. Missile defense systems, such as the Phalanx, currently installed on the surface fleet could be supplemented with a KV defense system. Tanks, infantry, fighting vehicles and armored personnel carriers are vulnerable to the wire-guided anti-tank missile. A KV defense system that intercepts these missiles may be utilized. 
     In all applications the divert thrusters cooperate with the attitude control thrusters to 1) position the reference axis (i.e. seeker axis or emitter axis) to a desired angular orientation relative to the threat, and 2) translate the body substantially perpendicular to the reference axis so as to eliminate the threat. The primary means of closure in either a decoy or kinetic kill is provided by the threat&#39;s velocity rather than the velocity of the KV. This obviates any requirement for a relatively powerful axial propulsion system which is required in guided missiles, such as the device disclosed in the Rae patent.