Patent Publication Number: US-10767968-B2

Title: Missile provided with a separable nose cone comprising at least one ejectable shell cooperating with a support element

Description:
The present invention concerns a missile provided with at least one ejectable shell forming part of a droppable or separable protective nose cone. 
     The present invention is applied, more particularly although not exclusively, to a missile comprising at least one propellant stage which is intended to propel the missile and which can be separated from the latter, as well as a terminal vehicle which is arranged at the front of this propellant stage and which carries out a terminal flight towards a target. Generally, such a terminal vehicle comprises at least one sensor forming part, for example, of a seeker, which is temperature-sensitive. 
     Although not exclusively, the present invention more specifically applies to a missile presenting a flying area remaining in the atmosphere and which has kinematic performances making it possible to bring the terminal vehicle to supersonic speeds. At these high speeds, the surface temperature of the missile can reach several hundred degrees Celsius under the effect of the aerothermal flow, which can be detrimental for the holding and the performances of the structures, electronic equipment and sensors present. Also, the missile is generally provided at the front with a protective nose cone, which generally comprises several individual shells, an which is intended to thermally and mechanically protect the terminal vehicle. 
     This protective nose cone, and at least some and preferably all of the shells, must be able to be removed at the right moment, in particular to make it possible to use the sensor placed on the terminal vehicle in the terminal phase of the flight. 
     Furthermore, in particular to ensure a good trajectory of the missile, the ejection angle of the shells must be controlled, i.e. the angle from which the shells of the nose cone are no longer connected to the body of the missile. 
     Different usual systems are known to eject the shells with the following problems. In particular:
         on subsonic missile flying in a low atmosphere, simply, generally, it is ensured that the shells of the nose cone do not close under the effect of the aerodynamic flow by guaranteeing a minimum opening angle. This is incompatible with a low-altitude and high-speed separation, as the shells would thus have a too important rotation speed and would risk to suddenly fall back on the body of the missile;   during the whole pre-decapping phase (transport logistics, flight, etc.), the nose cone is subjected to important loading factors likely to deform it. This is why usual articulation solutions do not make it possible to maintain the base of the nose cone; and   an architecture which provides that the shells of the protective nose cone are articulated on the terminal vehicle, generates a important residual mass on the vehicle, due in particular to the mass of hinges or articulations of the shells used for this purpose, and penalises its performances during the terminal flight, that is the most crucial phase.       

     These usual solutions are not satisfactory to make it possible for an ejection of at least one shell of a nose cone of the missile in the applications considered (for example, at a low altitude and at a high speed). 
     The present invention aims to overcome this disadvantage. It relates to a missile provided with a body presenting a longitudinal axis called main longitudinal axis and at least one separable nose cone, said nose cone comprising at least one ejectable shell, said shell being connected by a so-called rear end to a support element of the missile and being defined around a longitudinal axis called secondary longitudinal axis. 
     According to the invention, said support element presents a circular arc shape centred on the main longitudinal axis and arranged orthogonally in relation to the latter, said support element being provided with an edge assembly and a crown element each presenting a circular arc shape centred on the main longitudinal axis, said crown element being arranged coaxially inside said edge assembly so as to create a housing between them, the rear end of the shell presenting a thickness adapted to said housing so as to be able to be received in said housing with a transverse contact in the bottom of the housing, a first longitudinal contact with the edge assembly and a second longitudinal contact with the crown element, said edge assembly being designed so as to enable the shell to pivot in relation to the body of the missile from a mounting position wherein the secondary longitudinal axis of the shell is substantially parallel to said main longitudinal axis (preferably the secondary longitudinal axis of the shell is parallel by being combined with said main longitudinal axis) towards at least one pivoted position wherein the secondary longitudinal axis presents a non-zero angle in relation to said main longitudinal axis, said edge assembly also being designed so as to:
         maintain at least partially said first longitudinal contact with said rear end of the shell, while the shell presents an orientation in relation to the body of the missile for which said secondary longitudinal axis presents, in relation to said main longitudinal axis, an angle smaller than a predetermined angle, called ejection angle; and   end said first longitudinal contact with said rear end of the shell, as soon as said secondary longitudinal axis presents, in relation to said main longitudinal axis, an angle higher than or equal to said ejection angle.       

     Thus, thanks in particular to the configuration of said support element, combined with that of the rear end of the shell, it is possible to provide an angle (so-called ejection angle) from which the rear end of the shell is no longer radially in contact towards the outside (against said edge assembly), and the shell thus released from this contact (so-called first longitudinal contact) can be ejected from the missile, as specified below. These specific configurations and architectures, although particularly well-adapted to a missile flying at a low altitude and at a high speed, can be used on any type of missile, whatever its flying area. 
     Advantageously, said edge assembly comprises two circular arc edge sections, arranged symmetrically in relation to a longitudinal plane containing the main longitudinal axis, each of said edge section being designed such that its orthogonal projection over said longitudinal plane presents a rectilinear front edge forming with its rear edge an angle equal to said ejection angle, said rear edge being orthogonal to said main longitudinal axis. 
     In addition, advantageously, said rear end of the shell comprises, in thickness, a tapered rear portion intended to be received in a contacting manner in said housing, followed towards the front by a thick portion forming a shoulder making it possible for an auxiliary transverse contact of the shell on the front edge of the edge assembly in the mounting position. 
     In a first embodiment, said support element corresponds to a portion of the body of the missile. 
     Furthermore, in a second embodiment, said support element is an insert part, capable of being mounted on the body of the missile. 
     Preferably, the features, in particular thickness, are formed (preferably machined) directly in the rear end of the shell. However, in a specific embodiment, said rear end is provided with an interface portion which is fixed to the rear of the shell. 
     Furthermore, advantageously, the missile comprises at least one controllable actuation device, capable of generating a force likely to lead to a pivoting of the shell from the mounting position to an ejection position wherein the secondary longitudinal axis of the shell presents an angle equal to an ejection angle in relation to said main longitudinal axis of the body of the missile. 
     In a preferred embodiment, the missile comprises two complementary shells forming said nose cone, and an annular support part formed of two identical support elements, each of said shells being connected via its rear end to one of said support elements of the support part. 
    
    
     
       The appended figures will make it well understandable how the invention can be achieved. In these figures, identical references designate similar elements. 
         FIGS. 1 and 2  schematically show an example of missile to which applies the present invention, provided with a protective nose cone which is, respectively, in a mounted position on the missile and in an opening position. 
         FIG. 3  shows the nose cone in an opening position. 
         FIGS. 4 to 11  represent different schematic views showing the maintaining and ejecting of a nose cone shell in relation to the missile, these  FIGS. 4 to 11  making it possible to highlight well the main features of the invention. 
     
    
    
     The present invention applies itself to a missile  1  represented schematically in  FIGS. 1 and 2 . The missile  1  is provided with a body  7 , at least partially cylindrical, presenting a longitudinal axis X-X called main longitudinal axis. The missile  1  is provided at the front with a protective nose cone  2 . 
     This protective nose cone  2  (called “nose cone  2 ” below) comprises a plurality of shells  3  and  4 , in this case, two shells  3  and  4  in the examples considered in the description below. The adverbs “front” and “rear” are defined with respect to the movement direction F of the missile  1 . 
     In the specific example represented in  FIG. 1 , the missile  1  comprises at least one droppable propellant stage  5  (to the rear) and a terminal vehicle  6  which is arranged at the front (in the movement direction F) of this propellant stage  5 . 
     Generally, such a flying terminal vehicle  6  comprises, in particular, at least one sensor  8  arranged at the front, forming part, for example, of a seeker and likely to be temperature-sensitive. The propellant stage  5  and the terminal vehicle  6  which can be of any usual type, are not further discussed in the following description. Usually, the propellant stage(s)  5  of such a missile  1  are intended to propel said missile  1 , from the firing to the approach of a target (before being neutralised by the missile  1 ). The terminal phase of the flight is, itself, carried out autonomously by the terminal vehicle  6 , which uses, in particular, the information coming from the embedded sensor  8 , for example an optoelectronic sensor intended to assist with detecting the target. To do this, the terminal vehicle  6  comprises all the usual means (not further described), which are necessary to carry out this terminal flight. Before implementing the terminal phase, the nose cone  2  is dropped, after separation of the various shells  3  and  4 , by pivoting, as specified below, to release the (flying) terminal vehicle  6 , which is then separated from the remainder of the missile  1 . 
     The missile  1  is therefore provided at the front with a separable (or droppable) nose cone  2  which is intended, in particular, to thermally and mechanically protect the terminal vehicle  6 . This protective nose cone  2  must however be able to be removed at the right time, in particular to make it possible to use the sensor  8  placed on the terminal vehicle  6  in the terminal phase of the flight. 
     In the situation of  FIG. 1 , the nose cone  2  is mounted on the missile  1  in a so-called (protective) mounting position. The terminal vehicle  6  represented by a dashed line is mounted inside the nose cone  2 . 
     Furthermore, in the situation of  FIGS. 2 and 3 , the shells  3  and  4  are in the process of being separated, by being pivoted, as illustrated respectively by the arrows α 1  and α 2 , during an opening or dropping phase of the nose cone  2 . The release (or ejection) of the shells  3  and  4  and the impulse to generate the movements illustrated by the arrows α 1  and α 2  (deviating from the axis X-X), can be caused by a suitable actuation device  9 , for example a pyrotechnic actuator arranged preferably at the front of the nose cone  2  (inside the latter), as schematically represented in a dashed line in  FIG. 1 . 
     Although not exclusively, the present invention is particularly well-suited to a missile  1  having a flying area remaining in the atmosphere and which has kinematic performances making it possible to bring the terminal vehicle  6  to hypersonic speeds. At these high speeds, the surface temperature of the missile  1  can reach several hundred degrees Celsius under the effect of the aerothermal flow, which requires providing an effective nose cone  2  to make it possible for the stability and the performances of the structures, electronic equipment and embedded sensors. However, the present invention can be applied to a missile evolving in any case from the flying area (in and outside of the atmosphere) and for speeds going from the subsonic to the high supersonic/hypersonic. 
     According to the invention, the nose cone  2  is connected by a rear end  2 A to a support part  10  of the missile  1 , as represented in  FIG. 3 . In the example represented, the two shells  3  and  4  are connected, each, by the rear end  3 A and  4 A thereof to a support element  11 ,  12  ( FIGS. 4 and 7 ) forming part of the support part  10 . 
     Each of these shells  3  and  4  is defined around a longitudinal axis called secondary longitudinal axis L-L, as represented in  FIGS. 4 and 5 , in particular. 
     In the preferred embodiment, the annular support part  10  is formed of two identical support elements  11  and  12 . Each of the shells  3  and  4  is therefore connected to its rear end  3 A,  4 A to one of said support elements  11  and  12 . 
     In addition, according to the invention, each support element  11 ,  12  presents a circular arc shape centred on the main longitudinal axis X-X and arranged in a plane P ( FIG. 2 ) which is orthogonal to said axis X-X. 
     The controllable actuation device  9  is capable of generating a force (illustrated by a double arrow E in  FIGS. 2 and 3 ) able to lead to the shells  3  and  4  pivoting from the mounting position of  FIG. 1  to an ejection position, wherein the secondary longitudinal axis L-L of each shell  3 ,  4  presents an angle equal to a so-called ejection angle α 0  in relation to said main longitudinal axis X-X of the body  7  of the missile  1 , as illustrated in  FIG. 5  for the shell  3  and specified below. 
     As represented in  FIGS. 6 and 7 , each support element  11 ,  12  is provided with an edge assembly  13  and a crown element  14 . The edge assembly  13  and the crown element  14  present, each, a circular arc shape centred on the main longitudinal axis X-X. 
     In addition, the crown element  14  is arranged coaxially along the axis X-X, radially inside said edge assembly  13  so as to create, between them, a circular arc shaped housing  15 . 
     The embodiment of the invention below is described for the shell  3 . The embodiment is identical for the shell  4 . 
     The rear end  3 A of the shell  3  presents a thickness E 1  adapted to the radial gap of said housing  15  so as to be able to be received in said housing  15 . The rear end  3 A is received in the housing  15  (in the mounting position) preferably with a three-point contact, as represented in  FIG. 9B  which is an enlarged view of the portion V 1  of  FIG. 9A , namely:
         a transverse contact C 1  in the bottom  15 A of the housing  15 ;   a first longitudinal contact C 2  (radially external) with the edge assembly  13 ; and   a second longitudinal contact C 3  (radially internal) with the crown element  14 .       

     These contacts make it possible for a simple and effective maintenance of the shell  3  to its base (rear end  3 A). This maintenance is achieved from the integration of the shell until it is ejected. The longitudinal contacts C 2  and C 3  are not however always simultaneous and/or evenly distributed over the shell  3 . 
     In addition, the edge assembly  13  is designed so as to enable the shell  3  to pivot in relation to the body  7  of the missile  1 :
         from a mounting position (wherein the secondary longitudinal axis L-L of the shell  3  is substantially parallel to said main longitudinal axis X-X, preferably the secondary longitudinal axis L-L of the shell  3  is combined with the main longitudinal axis X-X), as represented in  FIG. 4 ;   towards at least one pivoted position (wherein the secondary longitudinal axis L-L presents a non-zero angle in relation to said main longitudinal axis X-X), as represented in  FIG. 5 .       

     In addition, the edge assembly  13  is also designed so as to:
         maintain (at least partially) said first longitudinal contact C 2  with said rear end  3 A of the shell  3 , while the shell  3  presents an orientation in relation to the body  7  of the missile  1  for which said secondary longitudinal axis L-L presents, in relation to said main longitudinal axis X-X, an angle smaller than said predetermined ejection angle α 0 ; and   to suppress said first longitudinal contact C 2  with said rear end  3 A of the shell  3 , as soon as said secondary longitudinal axis L-L presents in relation to said main longitudinal axis X-X, an angle higher than or equal to said ejection angle α 0 , as represented in  FIG. 5 .       

     Thus, thanks in particular to the configuration of said support element  11 ,  12  combined with that of the rear end  3 A,  4 A of the shell  3 ,  4 , there is an ejection angle α 0  from which the rear end  3 A,  4 A of the shell  3 ,  4  is no longer in contact radially towards the outside (against said edge assembly  13 ), and the shell  3 ,  4  thus released from this contact (called first longitudinal contact C 2 ) can be ejected from the missile  1 . 
     This configuration of the support element  11 ,  12  combined with that of the rear end  3 A,  3 B of a shell  3 ,  4 , or more generally the configuration of the support part  10  combined with that of the rear end  2 A of the nose cone  2 , forms a maintenance and ejection system S making it possible to maintain the nose cone  2  and making it possible for its ejection by controlling the ejection angle. 
     The ejection angle α 0  can, in particular, be adapted to the missile (type, size, etc.) considered and to the ejection conditions (altitude, atmosphere, trajectory of the missile, etc.) considered. This ejection angle α 0  can be refined by tests. Although not exclusively, the ejection angle α 0  can, for example, be defined in a value range from 6° to 15°. 
     As represented in  FIGS. 8 and 9A  in particular, said edge assembly  13  intended for a shell  3  or  4  comprises two circular arc edge sections  16 . These edge sections  16  are arranged symmetrically in relation to a longitudinal plane OXZ containing the main longitudinal axis X-X. 
     In  FIGS. 9A and 10A  in particular, a marker OXYZ has been represented, wherein 0 represents the intersection of the axis X-X with the plane P, OX is defined along the axis X-X in the direction F, OY is such that the plane OXY substantially corresponds to a separation plane between the shells  3  and  4 , and OZ is such that the plane OXZ substantially forms a symmetry plane for each of the shells  3  and  4 . 
     Each of said edge sections  16  is designed such that its orthogonal projection over said longitudinal plane OXZ presents a rectilinear front edge  17  forming with the (rectilinear) rear edge  18  thereof, an angle β equal to said ejection angle α 0 , as represented in  FIG. 11 . 
     Two edge sections  16 , one of which is intended for the shell  3  and the other to the shell  4  form, each time, an edge part  19 , as represented in  FIG. 11 . 
     The support part  10  therefore comprises two edge parts  19  of this type, which are systematically mounted in relation to the longitudinal plane OXZ, as shown in  FIG. 8 . In a specific embodiment, the two edge parts  19  are made of one (single) part of one single holding. 
     Likewise, the support part  10  comprises two crown elements  14 , identical and symmetrical in relation to the plane OXY. These two crown elements  14  form a crown  20  ( FIG. 7 ) centred on the axis X-X. This crown  20  is preferably an insert part. It can also correspond to a portion of the external surface of the terminal vehicle  6  as illustrated in  FIG. 8 . 
     In addition, the rear end  3 A of the shell  3  comprises, in thickness, a tapered rear portion  21  (of thickness E 1 ) intended to be received in a contacting manner in said housing  15 , followed towards the front of a thick portion  22  (of thickness E 2  higher than the thickness E 1 ) forming a shoulder  23  making it possible for an auxiliary transverse contact C 4  of the shell  3  on the front edge  17  of the edge assembly  13  in the mounting position, as represented in  FIG. 9B . This shoulder  23  presents a shape adapted to that of the front edges  17  of the two associated edge sections  16 . 
     Thus, as represented in  FIGS. 10A and 10B , at the level of the enlarged zone V 2  of  FIG. 10B  corresponding to the intersection of the axis OZ with the shell  3 , the rear end  3 A of the shell  3  does not comprise any tapered portion but only the thick portion  22  of thickness E 2 . 
     The system S makes it possible to maintain the shells  3  and  4  as illustrated by the arrows G in  FIG. 8 , and a pivoting of the shells  3  and  4  as illustrated by the arrows H in this  FIG. 8 . 
     In addition, the pivoting of the shell  3  is achieved without any hinges by simple contact at the level of a zone  25  ( FIGS. 2, 10A, 10B ) located in the proximity of the intersection of the axis OZ with the shell  3 . 
     In a first embodiment, the support part  10  corresponds to a portion of the body  7  of the missile  1 . 
     Furthermore, in a second embodiment, the support part  10  is an insert part, capable of being mounted (and fixed) on the body  7  of the missile  1 . 
     Moreover, preferably, the features, in particular of thickness (E 1  and E 2 ) are formed (preferably machined) directly in the rear end  3 A,  4 A of the shell  3 ,  4 . However, in an embodiment variant (not represented), the rear end  3 A,  4 A of each shell  3 ,  4 , presenting these features, is provided with an interface part which is fixed at the rear of the shell  3 ,  4 . 
     The functioning of the maintenance and ejection system S (controlling the ejection angle), such as described above, is as follows, during the ejection. 
     When the shells  3 ,  4  of the nose cone  2  must be separated, the actuation device  9  is activated to generate forces illustrated by the double arrow E (in  FIGS. 2 and 3 ) in order to make the shells  3  and  4  pivot in the directions illustrated by the arrows α 1  and α 2  ( FIG. 2 ). Thanks to the system S, the shells  3  and  4  are maintained on the support part  10  until the pivoting angles a 1  and α 2  reaching the value a 0  of ejection angle. In this pivoting position, the shells  3  and  4  are no longer maintained by the support part  10  and are released from the missile  1 , from which they deviate, which results in the dropping of the nose cone  2 . 
     The abovementioned features of the maintenance and ejection system S, and in particular the configuration of the support part  10  and the rear ends  3 A,  4 A of the shells  3 ,  4 , make it possible to control the separation angle of the shells  3  and  4  of the nose cone  2 . The ejection angle is an essential parameter which is difficult to control by usual solutions, according in particular to the ejection conditions (altitude, atmosphere, trajectory of the missile, etc.). Thanks to this controlling, it can be ensured that the ejection does not damage the missile and does not impede its terminal phase. 
     The system S functions in any case from the flying area (in and outside of the atmosphere) of a missile  1  and for speeds going from the subsonic to the high supersonic/hypersonic. 
     The system S thus presents numerous advantages. In particular:
         it is based on a purely mechanical architecture, which gives it an excellent repeatability;   it is based on a passive, simple, reliable and robust solution, which is adaptable to all types of missiles provided with ejectable (nose cone) shells;   the simplicity of the geometry minimises the mass embedded on the missile  1 , and guarantees its ease of production and integration;   in storage, logistical transport and flight phases, before de-capping, the system S makes it possible to regain forces between the shells  3  and  4 ; and   the architecture of the system S is fully configurable according to the flying area and for each of the shells  3  and  4  (with a possible asymmetry, if needed).