Abstract:
An airborne platform comprising: (a) an aerodynamic body; (b) a protected element within the aerodynamic body; and (c) a cover, reversibly secured to the aerodynamic body, for protecting the protected element from an external atmosphere. The present invention is particularly suited for an electro-optical detection system equipped with an optical dome or window. Optical windows carried on airborne platforms are known to be particularly sensitive to the high temperatures generated by air friction at super-sonic speeds. By covering such a window when the airborne platform is in flight, and releasing the cover just before use, the electro-optical detection system equipped with the optical window can be made operative, even at airborne platform speeds of several Mach and higher.

Description:
FIELD AND BACKGROUND OF THE INVENTION  
         [0001]    The present invention relates to a jettisonable element and a high speed missile utilizing same. More particularly, the present invention relates to a high speed missile including at least one jettisonable element that functions as a detachable cover protecting an optical window or dome from the external atmosphere, as a drag reduction element, and/or as a radar ghost when jettisoned.  
           [0002]    The navigation of a missile to target is achieved using a guidance system. One or more guidance systems are generally employed. Radar is one such guidance system. Although radar is effective, it is subject to interference, both intentional interference deployed as a defense mechanism, and accidental interference resulting from environmental conditions. Therefore, radar is often employed in conjunction with optical or electro-optical guidance systems, either of which may operate in the visible or infrared portion of the spectrum. These guidance systems are composed of a sensor or a detection system (e.g., electro-optical camera), and an analyzing system. The detection system must be onboard, although the analyzing system may be located outside the missile, for example at a base on the ground or in a platform such as an airplane which launched the missile, which communicates with the missile during flight. Alternatively, both the detection system and the analyzing system are carried on-board. This alternative, referred to as a “launch and forget” guidance system, is especially desirable in the case of missiles flying at high supersonic speeds where the time available for navigation decisions is extremely short, making communication with a remote location a practical impossibility.  
           [0003]    The detection system must have a sensor in communication with the environment. At the same time, the sensor must be protected from the environment. For optical or electro-optical guidance systems this protection typically takes the form of an optical window or dome. These windows or domes are transparent to transmissions in a chosen range of wavelengths, while being opaque to transmissions with a wavelength outside that range. These optical windows or domes are typically coated with a shielding material which gives the window or dome the desired optical properties. As explained by D. Harris in “Materials for Infrared Windows and Domes” (SPIE Optical Engineering Press, 1948), which is incorporated herein by reference, most common approaches to shielding include coating the optical window with an electrically conductive layer, covering the window with a metallic mesh, or increasing the conductivity of the material forming the window. In general, the thin electrically conductive coatings applied to the window are transparent at visible and/or infrared frequencies, but opaque to microwaves and radio waves. This makes such coatings useful in shielding sensitive electro-optical detectors against harmful electromagnetic interference (Kohin et al., SPIE Crit. Rev. CR 39:3-34(1992)). The shielding capabilities of these materials stems from their ability to reflect and/or absorb incident radiation. In general, the greater the conductivity of the coating material, the more effective the shielding. Common coating materials are described in, for example, (i) Pellicori and Colton (Thin Solid Films 209:109-115(1992)); (ii) Rudisill et al. (Appl. Opt. 13:2075-2080 (1974)), and (iii) Bui and Hassan (Proc. SPIE 3060:2-10(1997)), all of which are incorporated herein by reference. Since the conductivity of these materials decreases with increasing temperature, they lose their shielding effectiveness when they are heated. At the same time, transmission of desired wavelengths through the shield is often diminished by heating.  
           [0004]    The use of missiles that include electro-optical detection systems is often constrained to near-sonic speeds, because at very high speeds (e.g., above several Mach), friction from the air causes heating of the optical window or dome which protects the electro-optical detection system. This heating changes the conductivity of the coating on the optical window or dome and as such alters the optical properties thereof. This results in incapacitation of the detection system of the missile, either because transmissions in the chosen range of wavelengths no longer pass through the window or dome, or because interference (transmissions with a wavelength outside the chosen range) is allowed to pass through the window or dome.  
           [0005]    Consequently, prior art missiles flying at high speeds are substantially limited to radar guidance systems.  
           [0006]    There is thus a widely recognized need for, and it would be highly advantageous to have, a cover for protecting an optical window or dome of missiles (and airborne platforms in general) from an external environment, which cover can be jettisoned to allow target acquisition by the optical payload when necessary. It would be of further advantage if the cover would also serve to reduce drag. Such a cover can also serve as a radar reflecting element, and as such produce a radar ghost when jettisoned.  
         SUMMARY OF THE INVENTION  
         [0007]    According to the teachings of the present invention there is provided an airborne platform comprising: (a) an aerodynamic body; (b) a protected element within the aerodynamic body; and (c) a cover, reversibly secured to the aerodynamic body, for protecting the protected element from an external atmosphere.  
           [0008]    According to another aspect of the present invention there is provided an airborne platform comprising: (a) an aerodynamic body; (b) an electro-optical detection system situated within the aerodynamic body, the electro-optical detection system being equipped with an optical window; and (c) a cover, reversibly secured to the aerodynamic body, for protecting the optical window from an external atmosphere.  
           [0009]    According to further features in the described preferred embodiments, the airborne platform further includes: (d) a mechanism for at least partially detaching the cover from the aerodynamic body.  
           [0010]    According to still further features in the described preferred embodiments, the airborne platform further includes: (e) a securing assembly for securing the cover to the aerodynamic body.  
           [0011]    According to still further features in the described preferred embodiments, the cover is breakable, such that the releasing mechanism breaks the releasable element, thereby releasing the cover.  
           [0012]    According to still further features in the described preferred embodiments, the releasing mechanism is operative to first act against the releasable element to unsecure the second end of the cover, and only subsequently to act against aerodynamic force, thereby detaching the second end of the cover from the aerodynamic body.  
           [0013]    According to still further features in the described preferred embodiments, the securing assembly includes a hinge reversibly connecting a first end of the cover to a first region of the aerodynamic body, and a releasable element securing a second end of the cover to a second region of the aerodynamic body.  
           [0014]    According to still further features in the described preferred embodiments, the hinge is configured such that when the second end of the cover separates from the second region of the aerodynamic body when the airborne platform is in flight, an aerodynamic force exerted by the external atmosphere on the cover detaches the hinge, thereby removing the cover from the aerodynamic body.  
           [0015]    According to still further features in the described preferred embodiments, the hinge includes a stoppage element, the stoppage element serving for limiting an angular movement of the hinge such that when the second end of the cover separates from the second region according to a predetermined parameter, the force exerted on the cover breaks the cover in a predetermined location.  
           [0016]    As used herein and in the claims section that follows, the term “predetermined parameter” refers to a particular angle of rotation or a particular distance at which the cover and hinge structure is designed and made operative to detach the cover. In many hinges, the requisite separation of the cover from the second region for triggering the detachment of the cover can be defined either as an angle or as a distance (or some combination thereof).  
           [0017]    According to still further features in the described preferred embodiments, the hinge includes an asymmetric ball element disposed in a socket. Thus, when the second end of the cover separates from the second region by a predetermined angle, the asymmetric ball element is rotated until disengagement from the socket, effecting spontaneous disassembly of the ball element.  
           [0018]    According to still further features in the described preferred embodiments, this force includes an aerodynamic force exerted on the cover by the external atmosphere. According to still further features in the described preferred embodiments, this force includes a force delivered to the cover by the releasing mechanism.  
           [0019]    According to still further features in the described preferred embodiments, the hinge includes a shearable pin.  
           [0020]    According to still further features in the described preferred embodiments, the cover is operative to jettison away from the platform when released.  
           [0021]    According to still further features in the described preferred embodiments, the cover includes at least one radar reflective region, such that the jettisoning of the cover from the airborne platform generates a radar ghost.  
           [0022]    According to another aspect of the present invention there is provided a device for protecting an element in an airborne platform from an external atmosphere, the device comprising: (a) a cover, reversibly secured to an aerodynamic body of the airborne platform, for protecting the protected element from an external atmosphere; and (b) a mechanism for at least partially detaching the cover from the aerodynamic body.  
           [0023]    According to further features in the described preferred embodiments, the cover is operative to jettison away from the platform when released.  
           [0024]    According to still further features in the described preferred embodiments, the cover includes at least one radar reflective region, such that the jettisoning of the cover from the airborne platform generates a radar ghost.  
           [0025]    According to still further features in the described preferred embodiments, the securing assembly includes a hinge for connecting a first end of the cover to a first region of the aerodynamic body, and a releasable element securing a second end of the cover to a second region of the aerodynamic body.  
           [0026]    According to still further features in the described preferred embodiments, the device further includes: (e) a securing assembly for securing the cover to the aerodynamic body.  
           [0027]    According to still further features in the described preferred embodiments, the releasing mechanism includes a high pressure gas reservoir as a source of energy for a force acting upon the releasable element.  
           [0028]    According to still further features in the described preferred embodiments, the releasing mechanism includes a pyroelectric element as a source of energy for a a force acting upon the releasable element.  
           [0029]    According to still further features in the described preferred embodiments, the releasing mechanism is designed to first act against the releasable element to unsecure the second end of the cover, and only subsequently to act against the aerodynamic force, thereby detaching the second end of the cover from the aerodynamic body.  
           [0030]    According to another aspect of the present invention there is provided a method of protecting an element in an airborne platform, the airborne platform having an aerodynamic body, from an external atmosphere, the method comprising: (a) covering the element with a cover to the aerodynamic body; (b) securing the cover to the aerodynamic body; and (c) releasing and at least partially detaching the cover to expose the element.  
           [0031]    According to further features in the described preferred embodiments, the securing of the cover is achieved using a hinge for connecting a first end of the cover to a first region of the aerodynamic body, and a releasable element for securing a second end of the cover to a second region of the aerodynamic body.  
           [0032]    According to still further features in the described preferred embodiments, the releasable element is broken, thereby releasing the cover.  
           [0033]    According to still further features in the described preferred embodiments, the releasing of the cover and the at least partially detaching the cover are performed by a releasing mechanism in discrete, sequential stages.  
           [0034]    According to still further features in the described preferred embodiments, the releasing of the cover is performed in a first stage and the at least partially detaching of the cover is performed subsequently in a second stage, such that the releasing mechanism acts against aerodynamic force only in the second stage.  
           [0035]    The present invention successfully addresses the shortcomings of the existing technologies by providing a system for and method of protecting airborne elements such as optical domes and windows from the air friction and the high temperatures generated while flying at high speed. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0036]    The invention is herein described, by way of example only, with reference to the accompanying drawings. With specific reference now to the drawings in detail, it is stressed that the particulars shown are by way of example and for purposes of illustrative discussion of the preferred embodiments of the present invention only, and are presented in the cause of providing what is believed to be the most useful and readily understood description of the principles and conceptual aspects of the invention. In this regard, no attempt is made to show structural details of the invention in more detail than is necessary for a fundamental understanding of the invention, the description taken with the drawings making apparent to those skilled in the art how the several forms of the invention may be embodied in practice.  
         [0037]    In the drawings:  
         [0038]    [0038]FIG. 1 is a partly cross sectional partly side view of a missile including the jettisonable electro-optical window or dome cover according to the teachings of the present invention;  
         [0039]    [0039]FIG. 2 is a close up view of the cover region depicted in FIG. 1, showing the releasable securing assembly and releasing mechanism according to the present invention;  
         [0040]    [0040]FIG. 3 is a cross sectional view of one embodiment of a releasable hinge according to the present invention;  
         [0041]    [0041]FIG. 4 a - e  are cross sectional views of another embodiment of a breakable hinge according to sequence of events leading to hinge breakage;  
         [0042]    [0042]FIG. 5 is a cross sectional view of one embodiment of a shearable pin according to the present invention;  
         [0043]    [0043]FIG. 6 is a detailed cross sectional view of the releasing mechanism according to the present invention; and  
         [0044]    [0044]FIG. 7 is a perspective view of a breakable element utilized for securing the securing element according to the teachings of the present invention.  
     
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0045]    The present invention is of a jettisonable element which can be utilized as a cover for protecting an optical window or dome of a missile from the external atmosphere and/or as a radar ghost when jettisoned. Specifically, the present invention is of an element which can be utilized to shield an optical window or dome of a high speed missile from the external atmosphere when attached, and/or to generate a. radar ghost when jettisoned. According to the present invention, this element is detachably attached to a missile body via a securing assembly which enables jettisoning of the element from the missile when the missile is in flight.  
         [0046]    The principles and operation of the present invention may be better understood with reference to the drawings and accompanying descriptions.  
         [0047]    Before explaining at least one embodiment of the invention in detail, it is to be understood that the invention is not limited in its application to details of construction and the arrangement of the components set forth in the following description or illustrated in the drawings. The invention is capable of other embodiments or of being practiced or carried out in various ways. Also, it is to be understood that the phraseology and terminology employed herein is for the purpose of description and should not be regarded as limiting.  
         [0048]    The use of missiles that include electro-optical detection systems is often constrained to near-sonic speeds, because at very high speeds (e.g., above several Mach), friction from the air causes heating of the optical window or dome which protects the electro-optical detection system. This heating changes the conductivity of the coating on the optical window or dome and as such alters the optical properties thereof. This results in incapacitation of the detection system of the missile, either because necessary data (transmissions in the chosen range of wavelengths) no longer passes through the window or dome, and/or because interference data (transmissions with a wavelength outside the chosen range) is allowed to pass through the window or dome.  
         [0049]    Since electro-optical detection systems are typically utilized by missiles during final target acquisition stages, such systems are only deployed during the final stages of trajectory. Thus, as is further described herein, the present invention provides a jettisonable heat shield which is utilized in high speed missiles for shielding the optical window or dome from heat when the missile is in flight. This heat shield is provided with a releasable securing assembly such that during the final stages of trajectory the heat shield can be jettisoned to expose the optical window or dome to the external atmosphere such that target acquisition can be effected by the electro-optical detection system.  
         [0050]    As used herein and in the claims section which follows, the term “optical window” is a general term that also includes any type of optical dome.  
         [0051]    More generally, an optical window is one example of a protected element that may be housed within an airborne platform.  
         [0052]    As used herein and in the claims section which follows, the terms “airborne platform” and “missile” are used interchangeably to refer to any airborne vessel or projectile, including, but not limited to a launchable projectile carrying an explosive charge. Also included in the definition are self-propelled missiles and missiles which move primarily due to an initial force applied at launch. Further included in the definition are airplanes and the like (e.g., a pod suspended from the wing of an aircraft), particularly those flying at high speeds.  
         [0053]    Referring now to the drawings, FIG. 1 illustrates a missile capable of operating at high speeds, which is referred to herein as missile  10 .  
         [0054]    Missile  10  includes an aerodynamic body  12  which is provided with at least one flight control mechanism  13 , such as at least one flight control surface (fin), which serves for stabilizing missile  10  and directing it to a target.  
         [0055]    Aerodynamic body  12  serves for housing an electro-optical detection system  14  equipped with an optical window or dome  16 . Preferably, optical window or dome  16  is coated with an optical coating which is substantially transparent to radiation at the visible and/or the infrared portion of the electromagnetic spectrum and substantially opaque to radiation at the radio frequency and/or radar frequency portion of the electromagnetic spectrum.  
         [0056]    The optical coating is characterized by high conductivity. Examples of suitable optical coating include, but are not limited to, doped Gallium Arsenide coat and doped Germanium coat.  
         [0057]    Electro-optical detection system  14  typically includes one or more sensors, such as a Forward Looking Infrared (FLIR) or video camera, or any other focusing component provided with an array of photosensitive elements, e.g., a charge coupled device (CCD). The focusing component may include, for example, lenses, reflectors, beam splitters, mirrors, and prisms arranged or configured to direct and focus incident radiation to the array of photosensitive elements. Electro-optical detection system  14  may also include various electronic systems which control the sensors, analyze and interpret the signals received by the sensors, and control the final trajectory of missile  10  by maneuvering flight control mechanism  13 . Electro-optical detection system may also include means for receiving signals from outside of the missile and may also include means for transmitting signals from the missile. Such electro-optical guidance systems are well known in the art and as such no further description is given herein.  
         [0058]    Aerodynamic body  12  preferably also houses a guidance system  6  which serves to control flight path of missile  10  before approaching a final trajectory. Such guidance systems operate according to well known principles and typically utilize such technologies such as, but not limited to, radar guidance or satellite (GPS) guidance. Preferably, aerodynamic body  12  further includes a liquid or solid fuel propulsion system  18  which serves to propel missile  10  to high speeds. Such propulsion systems are well known in the art and as such, no further description is provided herein.  
         [0059]    Aerodynamic body  12  also includes a warhead  20 , which is designed to detonate prior to, during, or following impact of missile on target. Missile  10  further includes a cover  22  which is secured to aerodynamic body by a releasable securing assembly which is further described hereinbelow. According to this aspect of the present invention cover  22  is positioned and configured so as to cover and protect optical window or dome  16  from an external atmosphere. Thus, when the missile is operating at high speeds, cover  22  serves as a heat shield. The releasable securing assembly is configured so as to allow cover  22  to be controllably jettisoned from missile  10  when in flight, to thereby expose optical window or dome  16  to external atmosphere when approaching a target.  
         [0060]    According to another preferred embodiment of the present invention and as shown in FIG. 2, cover  22  is secured to aerodynamic body  12  via a releasable securing assembly  24 . Preferably, assembly  24  includes a hinge  26  for hingedly connecting a first end  28  of cover  22  to a first region  30  of aerodynamic body  12 . Hinge  26  can be any one of several types disclosed herein, but it will be appreciated that other types of hinges may be implemented by those skilled in the art. Assembly  24  also includes a securing element  32  for releasably securing a second end  34  of cover to a second region  36  of aerodynamic body  12 .  
         [0061]    As is further described hereinbelow, aerodynamic body further includes a releasing mechanism  38  for controllably unsecuring element  32 .  
         [0062]    As used herein and in the claims section which follows, the term “hinge” refers to a rotatable element. Various types of hinges are mentioned explicitly herein by way of example.  
         [0063]    The terms “release” and “unsecure” are interchangeably used herein to refer to the action of unlocking but not separating two or more components, whereas the terms “detach” and “separate” are used interchangeably to refer to physically separating, i.e., putting a distance between two or more components.  
         [0064]    According to another preferred embodiment of the present invention securing element  32  is a bolt. Bolt  32  preferably includes a threaded region which threads into a breakable element  33  attached to second region  36  of aerodynamic body  12  as is further described hereinbelow. It will be appreciated that although securing element  32  is exemplified herein as a bolt, it can be of any design capable of securing second end  34  of cover  22  to second region  36  of aerodynamic body  12 .  
         [0065]    In a preferred embodiment, bolt  32 , instead of being threaded into breakable element  33 , is secured to second region  36  by a shearable pin (an example of a shearable pin  52  is provided in FIG. 5).  
         [0066]    According to another preferred embodiment of the present invention, and as specifically shown in FIG. 3, hinge  26  is configured such that first end  28  of cover  22  detaches from first region  30  when second end  34  of cover  22  separates a predetermined distance from second region  36  of aerodynamic body  12  when missile  10  is in flight. To enable such detachment, hinge  26  can be of an asymmetric ball in socket configuration, by way of example. According to this configuration, when second end  34  of cover  22  separates from second region  36  of aerodynamic body  12  by a predetermined angle, ball element  40  moves in socket element  42  to a point where ball element  40  frees from socket  42 , thus enabling spontaneous disassembly of hinge  26  (i.e., ball element  40  falls out of socket element  42 ) and subsequent detachment of first end  28  of cover  22  from first region  30  when missile  10  is in flight.  
         [0067]    According to another preferred embodiment of the present invention, hinge  26  is a breakable hinge. According to this configuration, when second end  34  of cover  22  rotatably separates from second region  36  by a predetermined angle, a force exerted on cover  22  breaks hinge  26  at a structurally weakened region formed in hinge  26 . This force may be primarily an aerodynamic force applied by the external atmosphere when missile  10  is in flight, particularly at high speeds. Alternatively or additionally, the force may be delivered to cover  22  by releasing mechanism  38  (see FIGS. 2, 6).  
         [0068]    To force hinge  26  to break, region  30  of aerodynamic body  12  includes a stoppage element  44  which serves for limiting an angular movement of hinge  26 . Thus, when second end  34  of cover  22  separates from second region  36  a predetermined distance, hinge rotates to a stop against stoppage element  44 , following which, the aerodynamic force exerted by the external atmosphere on cover  22  when missile  10  is in flight, breaks hinge  26  at a designed weakened region thereof.  
         [0069]    According to another preferred embodiment of the present invention, and as specifically shown in FIGS. 4 a - c  the weakened region  50  is a region interconnecting hinge  26  to cover  22 . Region  50  can be structurally weakened by an introduction of a groove or by the use of structurally weaker material as compared to the material utilized to fabricate the regions of hinge  26  and cover  22  which surround region  50 . In any case, the design of region  50  ensures that the aerodynamic force exerted by the external atmosphere on cover  22 , when missile  10  is in flight breaks hinge  26  only at region  50 . The sequence of events which lead to this breakage are illustrated in FIGS. 4 a - e.    
         [0070]    It will be appreciated that this breakable hinge configuration can also be applied to an integral non-rotating hinge  26  in which case stoppage element  44  is not necessary.  
         [0071]    While reducing the present invention to practice, however, it was discovered that the above described hinge detachment configurations, although functional, suffer from inherent limitations which can lead to unwanted detachment. For example, the use of these configurations fails to provide an exact angle at which cover  22  totally disconnects from missile  10 . Precision in controlling this angle, however, is usually of great importance, because uncontrolled jettisoning may result in cover  22  striking and thereby destroying missile  10 . In the case of the weakened hinge, in particular, it is a practical impossibility to predict, in an exact manner, the hinge behavior under aerodynamic forces characterizing supersonic speeds. In the case of the ball and socket hinge, such behavior is also somewhat unpredictable because while the aerodynamic forces increase and the cover opens, the area of contact within the hinge arrangement (i.e., the contact between ball and socket surfaces) decreases.  
         [0072]    Thus, according to another and presently preferred embodiment of the present invention, and as specifically shown in FIG. 5, hinge  26  includes a shearable pin  52 . According to this configuration, hinge  26  rotates to a stop against stoppage element  44 , following which a force exerted on cover  22 , breaks shearable pin  52  to thereby detach region  28  of cover  22  from region  30  of aerodynamic body  12  and thereby disconnect cover  22  from missile  10 .  
         [0073]    The above-mentioned force may be primarily an aerodynamic force applied by the external atmosphere when missile  10  is in flight, particularly at high speeds. Alternatively or additionally, the force may be delivered to cover  22  by releasing mechanism  38  (see FIGS. 2, 6).  
         [0074]    As already mentioned hereinabove, and as shown in FIG. 2, aerodynamic body  12  includes a releasing mechanism  38  which serves to unsecure element  32  from breakable (or releasable) element  33 . Such unsecuring can be achieved via any one of several dedicated mechanism and configurations. For example, and as shown in FIG. 6, mechanism  38  includes a hydraulically, mechanically or pneumatically driven piston  60 , which when actuated, exerts a force of a predetermined magnitude on top of breakable element  33  to which securing element  32  is secured. This force breaks element  33 , thus releasing or unsecuring securing element  32 .  
         [0075]    Cover  22  is designed to substantially reduce the drag force acting upon missile  10  when cover  22  is secured thereto. Nonetheless, the aerodynamic force exerted by the external atmosphere on cover  22  of missile  10  is enormous when missile  10  achieves high supersonic speeds such that unsecuring securing element  32  per se is not necessarily sufficient for detaching cover  22 . Therefore, according to a preferred embodiment of the present invention, release mechanism  38  further serves to forcibly separate end  34  of cover  22  from region  36  of aerodynamic body  12  when missile  10  is in flight and thus under aerodynamic forces exerted by external atmosphere. This separation can be forcibly effected, for example, by piston  60  of the above described configuration of mechanism  38  following unsecuring of element  32 .  
         [0076]    In a preferred embodiment, releasing mechanism  38  is designed to first act against securing element  32  to unsecure end  34  of the cover  22  from region  36  of aerodynamic body  12 , and only subsequently to act against the aerodynamic force force and thereby detach end  34  of cover  22  from aerodynamic body  12 . By acting against the resistances (strength of breaking element  33  and resistance due to aerodynamic force) in sequentially, rather than in parallel, the required force generated by releasing mechanism  38  and exerted by piston  60  is significantly reduced.  
         [0077]    One presently preferred embodiment for accomplishing this is shown by way of example in FIG. 6, wherein breakable element  33  is separated from cover  22  by air gap  35 . When piston  60  exerts pressure on the top of breakable element  33 , air channel  35  provides space for breakable element  33  to give way, without forcing securing element  32  (or any other inner working) to distend beyond the form of aerodynamic body  12 , such that no aerodynamic force needs to be overcome at this stage. Subsequently, as securing element  32  is unsecured, and piston  60  forces securing element  32  out beyond the form of aerodynamic body  12 , only the magnitude of the aerodynamic force needs to be overcome.  
         [0078]    In a preferred embodiment, releasing mechanism  38  is actuated by a high-pressure gas reservoir as a source of energy for a force acting upon the releasable element.  
         [0079]    In another preferred embodiment, releasing mechanism  38  is actuated by a pyroelectric element as a source of energy for a force acting upon the releasable element.  
         [0080]    Several configurations of breakable element  33  can be realized by the present invention. For example, according to a preferred embodiment of the present invention and as specifically shown in FIG. 7, breakable element  33  includes a first region  62  for securing element  32 , and a second region  64  which is attached to region  62  in a manner which allows region  62  to break off from region  64  when a predetermined amount of pressure is applied to region  62 . A breakable configuration attaching regions  62  and  64  can be provided via the use of structural weakening, such as holes, or by using weaker material at the region of attachment.  
         [0081]    According to the present invention, cover  22 , releasable securing assembly  24  and releasing mechanism  38  are designed such that cover  22  can be jettisoned away from the missile in a manner which avoids collision therewith. In addition, missile  10  is designed such that following jettisoning of cover  22 , the aerodynamic properties and weight distribution of missile  10  are not substantially affected.  
         [0082]    Thus, the present invention provides a jettisonable cover which can be utilized to shield a window or dome of an electro-optical detection system from heat generated as a result of friction with the external atmosphere.  
         [0083]    According to another preferred embodiment of the present invention cover  22  can also serve as a radar ghost when jettisoned.  
         [0084]    Thus, according to this embodiment of the present invention, cover  22  preferably includes radar reflective regions. It will be appreciated that such radar reflective regions are preferably provided on an inside surface of cover  22  such that these regions are concealed from radar radiation when cover  22  is attached to missile  10 , and exposed to radar radiation only after cover  22  has jettisoned.  
         [0085]    It will further be appreciated that cover  22  can also be utilized as a radar ghost in subsonic or supersonic missiles which do not carry an electro-optical detection system or regardless of such systems.  
         [0086]    Thus, according to another aspect of the present invention, cover  22  can be utilized solely as a radar reflective element and as such can be configured of any shape, size or number and can be attached to any region of a missile. Preferably, in such cases, cover  22  is attached to a rearward section of a missile such that when jettisoned, the likelihood of collision between cover  22  and the missile is minimized. Although the various jettisonable configurations of the cover described above provides several unique advantages when incorporated into a high speed missile, it will be appreciated that other configurations can also be realized by the present invention.  
         [0087]    Although the invention has been described in conjunction with specific embodiments thereof, it is evident that many alternatives, modifications and variations will be apparent to those skilled in the art. Accordingly, it is intended to embrace all such alternatives, modifications and variations that fall within the spirit and broad scope of the appended claims. All publications cited herein are incorporated by reference in their entirety. Citation or identification of any reference in this application shall not be construed as an admission that such reference is available as prior art to the present invention.