• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    - Missile provided with a separable protective cap. The missile (1) comprises at least one separable propulsion stage (5) and a terminal vehicle (6) which is arranged in front of the separable propulsion stage (5), said missile (1) being provided with the before a separable protective cap (2) comprising at least two individual shells (3, 4), and a connecting piece (10A) connected to the missile (1), rearwardly beyond the position (P1) of the rear end (11) of the terminal vehicle (6), said protective cap (2) being configured for, when mounted on the missile (1), surround all said terminal vehicle (6) and being connected by a rear end to the connecting piece (10A) via articulated connecting elements (7).
    公开号:FR3022995A1
    申请号:FR1401421
    申请日:2014-06-25
    公开日:2016-01-01
    发明作者:Clement Quertelet
    申请人:MBDA France SAS；
    IPC主号:
    专利说明:

    [0001] The present invention relates to a missile provided with a releasable or separable protective cap. More specifically, the present invention applies to a missile comprising at least one propulsion stage which is intended to propel the missile and which can be separated from the latter, and a terminal vehicle which is arranged at the front of this stage. propulsive and which makes a terminal flight towards a target. Generally, such a terminal vehicle comprises at least one sensor, for example part of a homing device, which is sensitive to temperature. Although not exclusively, the present invention applies more particularly to a missile having a flight range remaining in the atmosphere and which has kinematic performance to bring the terminal vehicle at hypersonic 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 to the strength and performance of structures, and electronic equipment and sensors present. Also, the missile is generally provided at the front of a protective cap, which generally comprises several individual shells and which is intended to thermally and mechanically protect the vehicle terminal. This protection cap must be removed at the appropriate time, especially to allow the use of the sensor placed on the terminal vehicle in the terminal phase of the flight.
    [0002] A localized protection cap is often provided which is thus relatively light. But it is then necessary to thermally protect directly parts of the terminal vehicle that are not covered by the protective cap. The set is generally lighter but once the terminal vehicle disheveled, its agility is penalized by the mass of these elements of thermal protection.
    [0003] In particular, an architecture that provides that the hulls of the protective cover are articulated on the terminal vehicle generates a large residual mass on the vehicle, due in particular to the mass of hinges or hull joints used for this purpose, and penalizes its performance during the terminal flight. The present invention aims to overcome this disadvantage. It relates to a missile comprising at least one separable propulsion stage and a terminal vehicle which is arranged in front of the propulsion stage, said missile being provided at the front of a separable protective cover (or releasable) comprising at least two individual hulls. According to the invention, said missile comprises a connecting piece linked to said missile, rearward beyond the position of the rear end of the terminal vehicle, and said protective cap, when mounted on the missile, surrounds all said terminal vehicle and is connected by a rear end to the connecting piece by means of articulated connecting elements. Thus, thanks to the invention, there is provided a protection cap that is encompassing, that is to say, which completely surrounds the terminal vehicle in normal protection position. Such an encompassing protective cap is certainly larger and therefore heavier than a localized protection cap, but this bounding cap structure which is linked to the missile, rearward beyond the position of the rear end of the terminal vehicle (via the connecting piece), minimizes the residual mass on the terminal vehicle after the separation, as specified below. This minimization of the mass maximizes the performance of the terminal vehicle in the terminal phase (which is the most sensitive). Note that: - a localized protection cap is lighter than an enclosing protective cap as mentioned above, but requires to thermally protect all parts of the terminal vehicle that would not be covered by the protective cap. The set is generally lighter but, once the terminal vehicle disheveled, its agility is penalized by the entire mass of thermal protection become superfluous; and that a possible loss of performance of the missile in the first phase of the launch, with an over-all protection cap heavier than a localized protection cap, is compensated in particular by one or more powerful propulsion stages. Advantageously, said connecting piece has a general shape of a ring. In a first embodiment, said connecting piece is an intermediate part of the body of the missile, which is arranged between the terminal vehicle and the propulsion stage. Advantageously, this intermediate piece is able to be separated from said terminal vehicle. In a second embodiment, the protective cap, the connecting piece and the rotary connection elements (including hinges) form a one-piece assembly, the connecting piece being adapted to be fixed on a so-called support part of the missile. Preferably, this support part is an intermediate part of the body of the missile, which is arranged between the terminal vehicle and the propulsion stage, and which is capable of being separated from said terminal vehicle.
    [0004] Furthermore, in a particular embodiment, the missile comprises at least one unit for regulating the internal pressure. Advantageously, this internal pressure regulating unit comprises at least one valve arranged in at least one channel generating an air passage between the inside of the protective cap and the outside of the missile. Preferably, said at least one channel is formed in said intermediate piece. As due to the aerothermal flow (in the case of supersonic missiles for example) and the flight altitude, likely to be encountered by the missile, the pressure difference between the inside and the outside of the protective cap may be important, the internal pressure regulating unit prevents the cap from deforming in flight and creates an opening allowing the introduction of aerothermal flow likely to damage structures, equipment and a sensor of the vehicle terminal. In addition, advantageously, said intermediate piece is configured to support the terminal vehicle and comprises ejection elements of the latter. Furthermore, in a particular embodiment, the missile comprises intermediate support elements arranged between the protective cap and the end vehicle, these intermediate support elements being fixed on an inner face of the protective cap and simply being in contact with each other. an outer face of the terminal vehicle. Thus, thanks to this particular embodiment: - either, it is prevented that the terminal vehicle bends inside the protective cap; - Or, the terminal vehicle also participates in maintaining the protective cap, which ensures a reasonable sizing (sufficiently low mass) of the latter. In addition, advantageously, the missile also comprises at least one shear force recovery system between the shells of the protective cap. With this system, the hulls do not have to be too thick (and therefore too mass) to be able to benefit from sufficient stiffness. In addition, advantageously, the missile also comprises means configured to create electrical continuity between adjacent shells, electrically conductive, of the protective cap, which allows in particular to provide electromagnetic protection. The appended figures will make it clear how the invention can be realized. In these figures, identical references designate similar elements. Figures 1 and 2 show schematically an example of a missile to which the present invention applies, provided with a protective cap which is, respectively, in a position mounted on the missile and in a position of release or opening. Figures 3 and 4 schematically show a particular embodiment of the cap according to the present invention, respectively, in a mounting position and in a mounted position.
    [0005] Figures 5 and 6 schematically show an example of means of a shear force recovery system between the shells of the protective cap, respectively, over the entire protective cap and an enlarged portion of the protective cap. The present invention applies to a missile 1 diagrammatically shown in FIGS. 1 and 2, which is provided at the front (in the direction of movement F of said missile 1) with a protective cap 2. This protective cap 2 has a plurality of shells 3 and 4, in this case two shells 3 and 4 in the example of Figures 1 to 4. The missile 1 of longitudinal axis XX comprises at least one propulsion stage 5 releasable (at the rear ) and a terminal vehicle 6 which is arranged at the front (in the direction of movement F) of this propulsion stage 5. In general, such a flying terminal vehicle 6 comprises, in particular, at least one sensor 8 arranged at the front , for example part of a homing device and likely to be sensitive to temperature. The propulsion stage 5 and the terminal vehicle 6, which may be of any conventional type, are not described further in the following description. In the usual way, the propulsion stage or stages 5 of such a missile 1 are intended for the propulsion of said missile 1, from the firing until the approach of a target (to be neutralized by the missile 1). The terminal phase of the flight is, in turn, carried out autonomously by the terminal vehicle 6, which uses in particular the information from the onboard sensor 8, for example an optoelectronic sensor intended to assist in the detection of the target. To do this, the terminal vehicle 6 includes all the usual means (not further described), which are necessary to achieve this terminal flight. Before implementing the terminal phase, the protective cap 2 is jettisoned or at least open, after separation of the various shells 3 and 4, for example by pivoting, to release the terminal vehicle 6 (steering wheel) which then separates of the rest of the missile 1. The missile 1 is therefore provided at the front of a separable protection cap 2 which is intended, in particular, to thermally and mechanically protect the terminal vehicle 6. This protective cap 2 must however be able to be removed at the appropriate time, in particular to allow the use of the sensor 8 placed on the vehicle terminal 6 in the terminal phase of the flight. In the situation of Figure 1, the protective cap 2 is mounted on the missile 1 in an operating position (or protection). The vehicle terminal 6 is mounted inside the protective cap 2 which is shown in thick lines. In addition, in the situation of FIG. 2, the shells 3 and 4 are being separated, for example by being pivoted via rotary connection elements 7 shown diagrammatically in FIG. 2, as illustrated by arrows al and a2, during a phase of opening or release of the protective cap 2. The release of the shells 3 and 4 and the pulse to generate the movements illustrated by the arrows al and a2, can be generated by a suitable device 13, for example a pyrotechnic actuator preferably arranged in front of the cap 2 (inside the latter), as shown in Figures 1, 3 and 4. This phase of opening or release of the protective cap 2 allows the release of the terminal vehicle 6, which can for example be ejected out of the missile 1 by means of appropriate ejection means (not shown). Although not exclusively, the present invention can be applied more particularly to a missile 1 having a flight domain remaining in the atmosphere and which has kinematic performance to bring the vehicle terminal 6 at 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 the provision of a protective cap 2 effective to allow the holding and performance of the structures, electronic equipment and embedded sensors. According to the invention, said missile 1 comprises a connection piece 10A, 10B linked to the missile 1, towards the rear (in the direction opposite to the direction of movement F) beyond the position P1 of the rear end 11 of the terminal vehicle 6 when mounted on the missile 1. In addition, according to the invention, when it is mounted on the missile 1, the protective cap 2 surrounds all said terminal vehicle 6 and is linked by a rear end 12 to the connecting piece 10A, 10B by means of articulated connecting elements 7, in particular hinges or other usual rotating elements. The protection provided by the protective cap 2 thus benefits not only the sensor 8, but also the entire terminal vehicle 6. The protective cap 2 encompasses the entire vehicle terminal 6 and is removed just before the use of the sensor 8 and the autonomous flight of the terminal vehicle 6. The autonomous flight time of the vehicle terminal 6 (with a use of the sensor 8 ) being short, it can indeed do without thermal protection during the terminal phase of the flight. Thus, thanks to this protective covering cap 2, which is removed before the autonomous flight of the terminal vehicle 6, the mass linked to the protection function (necessary only before this autonomous flight) is not impinged on the terminal vehicle 6. Said connecting piece 10A has a general shape of ring, whose outer diameter is substantially equal to the diameter of the body of the missile 1 in the part where is provided this connecting piece 10A.
    [0006] In a first embodiment shown in FIG. 1, the connecting piece 10A is an intermediate piece 15 of the body of the missile 1, which is arranged between the end vehicle 6 and the propulsion stage 5. This intermediate piece 15 is suitable for be separated from said terminal vehicle 6. The shells 3 and 4 of the protective cap 2 are thus articulated on the intermediate part 15 and the associated connecting means, in particular the rotary connection elements 7, are secured to this intermediate part 15 which can separate from the vehicle terminal 6 before the autonomous flight of the latter. This embodiment makes it possible in particular: an industrial division between the different subsystems (protective cover 2, terminal vehicle 6, intermediate part 15, and propulsion stage (s) 5); supporting the terminal vehicle 6 and integrating ejection devices (not shown) of the latter; and integrating an internal pressure regulating unit, specified below, away from the aerothermal flow (i.e., away from the nose 27 of the protective cap 2), for greater efficiency. In a second embodiment (shown in FIGS. 3 and 4), the protective cap 2, the connecting piece 10B (in the form of a ring or ring) and the rotary connection elements 7 form a one-piece assembly 16. clearly specify this one-piece assembly 16, it is represented: in a mounting position in FIG. 3, the assembly 16 being displaced rearward in the direction E, coaxially with the axis XX, until its rear end 12 arrives in the right position. It is then fixed on the missile 1; and in a position mounted in FIG. 4. In this mounted position, the connecting piece 10B is fixed on support means 17 of a support part 18 of the missile 1, via appropriate fixing means 19. Any type of support means 17 and fastening means 19, usual and cooperating, likely to achieve satisfactory attachment of the assembly 16 on the missile 1, may be considered. Preferably, the support portion 18 is an intermediate part of the body of the missile 1, which is arranged between the terminal vehicle 6 and the propulsion stage 5, for example similarly to the intermediate part 15 of the first embodiment mentioned above.
    [0007] This second embodiment makes it easier to manufacture and integrate the protective cap 2. In addition, by adapting the connecting piece 10B and possibly the fixing means 19, it is possible to easily adapt the assembly. 16 to different types of existing missile.
    [0008] Furthermore, in a particular embodiment, the missile 1 comprises at least one unit 20 for regulating the internal pressure. As shown diagrammatically in FIG. 1, this internal pressure regulation unit 20 comprises at least one channel 21 creating an air passage between the inside 22 of the protective cap 2 and the outside 23 of the missile 1, and at least one valve 24 which is arranged in said channel 21. In a particular embodiment, the channel or channels 21 are formed in said intermediate piece 15 as shown in FIG. 1, or in the intermediate piece 18 of FIGS. Thus, the internal pressure regulating unit 20 is arranged far from the aerothermal flow (i.e., away from the nose 27 of the protective cap 2), thereby increasing efficiency. The valve 24 may, for example, consist of a ball and a return spring thereof, sized so that the internal pressure in the protective cap 2 never exceeds a predetermined threshold (for example a few millibars ). Other usual embodiments of valve architecture can be used. As because of the aerothermal flow (in the case of supersonic missiles for example) and the flight altitude, likely to be encountered by the missile 1, the pressure difference between the inside 22 and the outside 23 of the protection cap 2 may be important, the unit 20 for regulating the internal pressure makes it possible to prevent the protective cap 2 from deforming in flight and creates an opening allowing the aerothermal flow to be introduced, which may damage the structures. , the equipment and in particular the sensor 8 of the terminal vehicle 6.
    [0009] Therefore, in such an embodiment, as shown in FIG. 1, the intermediate piece 15 interfaces with the propulsion stage 5 and the junction with the terminal vehicle 6, and serves as a passage for the channel 21 and also in a particular embodiment, the intermediate piece 15, 18 is configured to support the end vehicle 6, and is provided with conventional ejection elements (not shown) for ejecting this last. Moreover, in a particular embodiment, the missile 1 comprises intermediate support elements 26 which are arranged between the protective cap 2 and the end vehicle 6 in the mounted position of FIGS. 1 and 4. These intermediate support elements 26 are: - on the one hand, fixed (by an end 26A) on an inner face 2A of the protective cap 2, as shown in Figure 1; and - secondly, simply in contact (by the other end 26B) of an outer face 6A of the vehicle terminal 6, for example via a sole or a suitable pad.
    [0010] Thus, thanks to this particular embodiment, the terminal vehicle 6 also participates in maintaining the protective cap 2, which ensures a reasonable dimensioning (sufficiently low mass) of the latter. With this particular embodiment, in an alternative embodiment, it can provide a significant stiffness of the protective cap 2 so as to prevent, with the aid of the intermediate support elements, the terminal vehicle 6 (having in particular a large dimension) bends inside the protective cap 2. In the second embodiment of Figures 3 and 4, these intermediate support members 26 are part of the one-piece assembly 16. Furthermore, the missile 1 also comprises at least one system 28 for taking up shear forces between the shells 3 and 4 of the protective cap 2, as shown in FIGS. 5 and 6. This system 28 makes it possible to take up the shearing forces between the shells 3 and 4 , which do not have to be thick (and therefore too much mass) to benefit from sufficient stiffness.
    [0011] In the particular embodiment (given by way of example) of FIGS. 5 and 6, this system 28 comprises a plurality of connection positions 29 distributed along the junction between the two shells 3 and 4. Each of these positions of link 29 comprises: - an oblong recess 30 formed in a shell 4 along its wall; and a pin 31 which is fixed on the other shell 3, and which is movable in the oblong recess 30 along the wall, but which prevents transverse movement. In the context of the present invention, other types of junctions are possible between the shells 3 and 4 of the protective part 2. In particular, it is possible to envisage an internal covering, with edges of cooperating shapes or with a connection of type tenon / mortise, over the entire periphery of the junction or a large part of it. Furthermore, in a particular embodiment, the shells 3 and 4 of the protective cap 2 are electrically conductive, either by being made of an electrically conductive material, or by including electrical conduction means. Many different means are possible for this purpose, such as a metal film or a metal braid which covers a structural part of each of the shells.
    [0012] In this particular embodiment, the missile 1 also comprises means for ensuring electrical continuity between the electrically conductive shells 3 and 4 of the protective cap 2. These means may comprise, as represented by way of example in the figure 6, a seal 32, in particular a charged elastomer or a metal braid, which is arranged at the junction between the two shells 3 and 4 so as to provide electrical continuity. In the context of the present invention, other variants of embodiment are also possible to ensure electrical continuity. In particular, it is possible to envisage an electrically conductive element (or plate), which connects two caps inwardly covering the junction.
    [0013] This particular embodiment makes it possible to avoid the generation of electric arcs at the junction and to provide electromagnetic protection. Thanks to the invention, there is therefore provided a protective cap 2 which is encompassing, that is to say which completely surrounds the vehicle terminal 6 in normal protection position. Such a protective covering cap 2 is certainly heavier than a localized protection cap, but it minimizes the residual mass on the terminal vehicle 6 after the separation, since the means 7, 26 for protecting and articulating the shells 3 and 4 are not attached to the vehicle terminal 6 but dropped elements. This minimization of the mass maximizes the performance of the terminal vehicle 6 in the terminal phase (the most sensitive). Note that a possible loss of performance of the missile 1 in the first phase of the launch, with a protection cap 2 encompassing heavier than a localized protection cap, is compensable, including providing one or more powerful propulsion stages 5 . The overarching architecture of the protective cap 2, as described above, also has (with respect to a more localized protective cap), the following advantages: - increased protection; and - a greater flexibility to the developments of realization of the terminal vehicle 6 and / or the propulsion stage 5.
    权利要求:
    Claims (13)
    [0001]
    REVENDICATIONS1. Missile comprising at least one separable propulsion stage (5) and a terminal vehicle (6) which is arranged in front of the separable propulsion stage (5), said missile (1) being provided at the front of a cap separable protection device (2) comprising at least two individual shells (3, 4), characterized in that it comprises a connecting piece (10A, 10B) connected to the missile (1), rearwardly beyond the position (P1) of the rear end (11) of the terminal vehicle (6), and in that said protective cap (2), when mounted on the missile (1), surrounds all said terminal vehicle (6). ) and is connected by a rear end to the connecting piece (10A, 10B) via articulated connecting elements (7).
    [0002]
    2. Missile according to claim 1, characterized in that said connecting piece (10A, 10B) has a general shape of a ring.
    [0003]
    3. Missile according to one of claims 1 and 2, characterized in that said connecting piece (10A) is an intermediate piece (15) of the body of the missile (1), which is arranged between the terminal vehicle (6) and the propulsion stage (5).
    [0004]
    4. Missile according to claim 3, characterized in that this intermediate piece (15) is adapted to be separated from said terminal vehicle (6).
    [0005]
    5. Missile according to one of claims 1 and 2, characterized in that said protective cap (2), said connecting piece (10B) and said rotary connecting elements (7) form a one-piece assembly (16), said connecting piece (10B) being adapted to be fixed on a said support portion (18) of the missile (1).
    [0006]
    6. Missile according to claim 5, characterized in that this support portion (18) is an intermediate part of the body of the missile (1), which is arranged between the terminal vehicle (6) and the propulsion stage (5).
    [0007]
    7. Missile according to any one of the preceding claims, characterized in that it comprises at least one unit (20) for regulating the internal pressure.
    [0008]
    8. Missile according to claim 7, characterized in that said unit (20) for regulating the internal pressure comprises at least one valve (24) arranged in at least one channel (21) generating an air passage between the interior (22) of the protective cap (2) and the outside (23) of the missile (1).
    [0009]
    9. Missile according to one of claims 3 and 6, and claim 8, characterized in that said at least one channel (21) is formed in said intermediate piece (15, 18).
    [0010]
    10. Missile according to one of claims 3 and 6, characterized in that said intermediate piece (15, 18) is configured to support the terminal vehicle (6) and comprises ejection elements of the latter.
    [0011]
    11. Missile according to any one of the preceding claims, characterized in that it comprises intermediate support elements (26) arranged between the protective cap (2) and the terminal vehicle (6), said intermediate support elements ( 26) being fixed on an inner face (2A) of the protective cap (2) and being in contact with an outer face (6A) of the terminal vehicle (6).
    [0012]
    12. Missile according to any one of the preceding claims, characterized in that it comprises at least one system (28) for taking up shear forces between the shells (3, 4) of the protective cap (2).
    [0013]
    13. Missile according to any one of the preceding claims, characterized in that it comprises means (32) configured to create electrical continuity between adjacent shells (3, 4), electrically conductive, the protective cap (2). ).
    类似技术:
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    同族专利:
    公开号 | 公开日
    EP2960618A1|2015-12-30|
    PL2960618T3|2017-09-29|
    WO2015197922A1|2015-12-30|
    EP2960618B1|2017-05-10|
    IL249531A|2020-05-31|
    FR3022995B1|2017-06-09|
    JP6548678B2|2019-07-24|
    US20170131076A1|2017-05-11|
    IL249531D0|2017-02-28|
    JP2017519177A|2017-07-13|
    ES2628256T3|2017-08-02|
    US10054411B2|2018-08-21|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US4699062A|1987-02-11|1987-10-13|The United States Of America As Represented By The Secretary Of The Navy|VLA airframe clamshell opener assembly|
    EP1013546A2|1998-12-24|2000-06-28|National Space Development Agency Of Japan|Rocket payload fairing and method for opening same|
    WO2001034468A1|1999-11-11|2001-05-17|Honeywell Inc.|Method and apparatus for minimizing the transmission of acoustic vibrations to a payload|
    US3446464A|1967-03-09|1969-05-27|William A Donald|Method and apparatus for reducing sonic waves and aerodynamic drag|
    US3706281A|1971-04-01|1972-12-19|Nasa|Method and system for ejecting fairing sections from a rocket vehicle|
    DE2607336C2|1976-02-23|1983-12-22|Rheinmetall GmbH, 4000 Düsseldorf|Carrier projectile for ejectable bodies|
    JPS52162700U|1976-06-02|1977-12-09|
    JPS62103299A|1985-10-31|1987-05-13|Nissan Motor|Head drum section of rocket|
    US4727451A|1987-01-06|1988-02-23|The Boeing Company|Metallic conduction method and system for joined sections of composite structures|
    US5529264A|1994-02-18|1996-06-25|Lockheed Missiles & Space Company, Inc.|Launch vehicle system|
    RU2072097C1|1994-09-22|1997-01-20|Борис Константинович Бессчетнов|Method of separation of ballistic missile nose cone|
    JP3770430B2|1997-06-30|2006-04-26|株式会社アイ・エイチ・アイ・エアロスペース|Nose fairing separation device for flying objects|
    US6247671B1|1998-09-23|2001-06-19|Accurate Automation Corporation|Ion doping apparatus and method for aerodynamic flow control|
    US6193187B1|1998-12-31|2001-02-27|Harry Scott|Payload carry and launch system|
    US7082878B2|2003-07-01|2006-08-01|Raytheon Company|Missile with multiple nosecones|
    KR100758855B1|2006-06-02|2007-09-19|국방과학연구소|Apparatus for opening missile airframes by pyrotechnical power|
    US9329007B2|2013-02-01|2016-05-03|Orbital Atk, Inc.|Charged projectiles and related assemblies, systems and methods|
    US9534563B2|2013-12-09|2017-01-03|Raytheon Company|Cluster rocket motor boosters|
    JP2015223963A|2014-05-28|2015-12-14|株式会社Ｉｈｉエアロスペース|Rocket nose cone opening system|
    FR3022885B1|2014-06-25|2016-10-21|Mbda France|STRUCTURING WALL OF MISSILE, ESPECIALLY FOR THERMAL PROTECTION COFFEE|FR3065521B1|2017-04-21|2019-06-28|Mbda France|MISSILE PROVIDED WITH A SEPARABLE HEADBOARD COMPRISING AT LEAST ONE EJECTABLE SHELL COOPERATING WITH A SUPPORTING ELEMENT|
    CN109494469A|2018-07-13|2019-03-19|中国航空工业集团公司济南特种结构研究所|A kind of revolving body antenna house root error protection structure|
    CN109625338A|2018-12-12|2019-04-16|湖北航天飞行器研究所|The radome fairing and rocket that can be cast aside certainly|
    CN110749250B|2019-10-30|2021-03-16|中国科学院力学研究所|Hypersonic aircraft safety cover|
    CN111392010B|2020-04-06|2021-09-17|西北工业大学|Asymmetric buffering head cap for high-speed underwater entry of aircraft|
    CN111391992A|2020-04-06|2020-07-10|西北工业大学|High-speed underwater vehicle entering buffering head cap|
    US11220358B1|2020-08-21|2022-01-11|Brandon West|Hypersonic harmonic vehicle exciter and methods of use thereof|
    CN112413039B|2020-11-19|2021-10-22|大连理工大学|Damping type sailing body water-entry buffering device|
    法律状态:
    2015-06-30| PLFP| Fee payment|Year of fee payment: 2 |
    2016-01-01| PLSC| Search report ready|Effective date: 20160101 |
    2016-07-08| PLFP| Fee payment|Year of fee payment: 3 |
    2017-06-30| PLFP| Fee payment|Year of fee payment: 4 |
    2018-06-27| PLFP| Fee payment|Year of fee payment: 5 |
    2019-12-26| PLFP| Fee payment|Year of fee payment: 6 |
    2020-06-30| PLFP| Fee payment|Year of fee payment: 7 |
    2021-06-30| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1401421A|FR3022995B1|2014-06-25|2014-06-25|MISSILE PROVIDED WITH A SEPARABLE PROTECTIVE VEST|FR1401421A| FR3022995B1|2014-06-25|2014-06-25|MISSILE PROVIDED WITH A SEPARABLE PROTECTIVE VEST|
    EP15290152.6A| EP2960618B1|2014-06-25|2015-06-10|Missile with removable protective cap|
    PCT/FR2015/000114| WO2015197922A1|2014-06-25|2015-06-10|Missile provided with a separable protective fairing|
    US15/318,378| US10054411B2|2014-06-25|2015-06-10|Missile provided with a separable protective fairing|
    ES15290152.6T| ES2628256T3|2014-06-25|2015-06-10|Missile provided with a separable protection shield|
    JP2016573559A| JP6548678B2|2014-06-25|2015-06-10|Missile with separable protection fairing|
    PL15290152T| PL2960618T3|2014-06-25|2015-06-10|Missile with removable protective cap|
    IL249531A| IL249531A|2014-06-25|2016-12-13|Missile provided with a separable protective fairing|
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