EP3396300B1 - Dispositif d'actionnement pour l'éjection d'au moins une partie amovible de missile, en particulier d'une coiffe - Google Patents

Dispositif d'actionnement pour l'éjection d'au moins une partie amovible de missile, en particulier d'une coiffe Download PDF

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Publication number
EP3396300B1
EP3396300B1 EP18290030.8A EP18290030A EP3396300B1 EP 3396300 B1 EP3396300 B1 EP 3396300B1 EP 18290030 A EP18290030 A EP 18290030A EP 3396300 B1 EP3396300 B1 EP 3396300B1
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EP
European Patent Office
Prior art keywords
missile
piston
retaining rod
pyrotechnic
retaining
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
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Application number
EP18290030.8A
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German (de)
English (en)
French (fr)
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EP3396300A1 (fr
Inventor
Clément Quertelet
Clyde Laheyne
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
MBDA France SAS
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MBDA France SAS
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Publication date
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Priority to PL18290030T priority Critical patent/PL3396300T3/pl
Publication of EP3396300A1 publication Critical patent/EP3396300A1/fr
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Publication of EP3396300B1 publication Critical patent/EP3396300B1/fr
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/36—Means for interconnecting rocket-motor and body section; Multi-stage connectors; Disconnecting means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/34—Protection against overheating or radiation, e.g. heat shields; Additional cooling arrangements

Definitions

  • the present invention relates to an actuation device allowing the ejection of at least one removable part of a missile, and a missile provided with at least one such actuation device.
  • the present invention can be applied to a missile comprising at least one releasable propulsion stage and a terminal vehicle which is arranged at the front of the propulsion stage.
  • a terminal vehicle generally comprises, in particular, a sensor forming, for example, part of a seeker and capable of being sensitive to temperature.
  • the present invention can be applied to a missile having a flight domain remaining in the atmosphere and which has kinematic performance such that the terminal vehicle can be brought to 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 for the behavior and performance of the structures, electronic equipment and sensors present.
  • a cap generally comprising several individual shells, is arranged at the front of the missile, so as to thermally and mechanically protect the terminal vehicle during the flight phase of the missile. The cap is then ejected at the appropriate time to allow, in particular, the use of the sensor arranged on the terminal vehicle, during the terminal phase of the flight.
  • the cap ejection is implemented by an actuating device configured to generate sufficient force to separate the individual hulls in a very short time in order to make the sensor quickly operational and to avoid any disturbance of the missile performance during of the cap ejection phase.
  • the actuating device must take into account the thermal and mechanical constraints to which the individual hulls are subjected before the terminal flight phase.
  • One solution could consist in using a pyrotechnic actuator such as a pyrotechnic ejector bolt, to generate the force necessary for the separation of the individual hulls in very short times.
  • a pyrotechnic actuator such as a pyrotechnic ejector bolt
  • the temperatures of several hundred degrees Celsius to which the individual hulls are subjected risk degrading the operation of the pyrotechnic actuator fixed to them, or even triggering it inadvertently.
  • the products ejected and the blast effect of the pyrotechnic reaction are liable to damage the sensor of the terminal vehicle or to obstruct its measurement capacity, for example by depositing powder residues. This solution is therefore not applicable.
  • WO 2009/095910 describes a device for separating a missile cap.
  • the present invention aims to remedy these drawbacks. It relates to an actuating device allowing the ejection of at least one removable part of a missile, in particular at least one individual shell of a cap.
  • said pyrotechnic actuator is configured to be able to generate a force capable of breaking said at least one retaining rod.
  • a first end of said at least one retaining rod and one end of said pyrotechnic actuator are intended to be fixed to a missile element and a second end, opposite to said first end of said at least one holding rod, is intended to be fixed to said removable part of the missile.
  • an actuation device for the ejection of a removable missile part, such as an individual shell of a cap, which comprises a pyrotechnic actuator whose operation is made compatible with the thermal and mechanical constraints of the missile by the arrangement of at least one thermal insulation element and at least one retaining rod.
  • the pyrotechnic charge which is an element of the pyrotechnic actuator sensitive to the high temperatures to which the individual shells are subjected, is isolated from the heat flows in the cap by the arrangement of at least one thermal insulation element.
  • this localized thermal protection makes it possible to minimize the mass and the size of the on-board actuation device.
  • the actuation device guarantees mechanical maintenance during the flight phase.
  • the pyrotechnic actuator being only fixed to the removable part, preferably a cap shell, by one of its ends, the device actuator is provided with one or more retaining rods which provide the mechanical connection between this removable part and a fixing element, for example two individual shells of a cap.
  • these retaining rods are configured to support in particular the mechanical stresses of the cap during the phase of flight preceding the ejection of the cap.
  • these retaining rods comprise at least one part integral with said pyrotechnic actuator by means of a mechanical yoke, which ensures, for example, better stability of the device against mechanical stresses during the flight phase of the missile. and ejecting the cap.
  • said at least one retaining rod has a weakening zone, which is preferably located near the free end of the piston.
  • said at least one retaining rod is provided with at least one retaining element, located at the level of the mechanical yoke.
  • This retaining element is advantageously arranged to prevent any translational movement of said at least one retaining rod relative to the pyrotechnic actuator.
  • said at least one retaining rod is provided with at least one thermal insulation sleeve, at least on a section of the latter.
  • Said at least one thermal insulation sleeve is preferably located at the level of the mechanical screed. The advantageous arrangement of said at least one sleeve contributes to the thermal insulation of said pyrotechnic actuator.
  • said thermal insulation elements can be made of a material of the mica, mullite, or muscovite type.
  • the second end of said retaining rod is advantageously provided with a thread, arranged to allow the fixing of said retaining rod to a solid element of the removable part of the missile by means of a nut.
  • the present invention also relates to a missile which is provided with an actuating device such as that described above, said actuating device being fixed by a first end to a fastening element of a first part of the missile, by example an individual shell of a cap or a fixed element of the structure of the missile and by a second end, opposite the first end, to an element for fixing a removable part of the missile.
  • an actuating device such as that described above, said actuating device being fixed by a first end to a fastening element of a first part of the missile, by example an individual shell of a cap or a fixed element of the structure of the missile and by a second end, opposite the first end, to an element for fixing a removable part of the missile.
  • this removable part can correspond to any element to be ejected from the missile during its flight, and preferably to an individual shell of a cap.
  • said missile is provided with a cap comprising at least two individual shells, said first part represents one of said individual shells and said second removable part represents the other individual shell.
  • the actuation device is configured to separate and simultaneously separate the two individual hulls in order to eject them from the missile.
  • At least one thermal insulation element is advantageously fixed to a fixing element of at least one of said removable parts of the missile, and arranged opposite the free end of said piston.
  • the present invention applies to a missile 1 shown diagrammatically on the Figures 1 and 2 , which is provided at the front (in the direction of movement F of said missile 1) with a (protective) cap 2 comprising several removable parts, in this case a plurality of shells 3, 4.
  • the present invention relates to a actuating device 7 for ejecting the cap 2.
  • the present invention can be applied to any type of missile 1 comprising at least one removable part to be ejected.
  • the missile 1 with a longitudinal axis LL comprises at least one releasable propulsion stage 5 and a terminal vehicle 6 which is arranged in front of this propulsion stage 5.
  • such a flying terminal vehicle 6 comprises, in particular, at least one sensor 8 arranged upstream, for example forming part of a seeker and capable of being sensitive to temperature.
  • the propulsion stage 5 and the terminal vehicle 6 which can be of any usual type, are not described further in the following description.
  • the propulsion stage or stages 5 of such a missile 1 are intended for the propulsion of said missile 1, from firing to approaching a target (to be neutralized by missile 1).
  • the terminal phase of the flight is, in turn, carried out autonomously by the terminal vehicle 6, which in particular uses information from the on-board sensor 8, for example an optoelectronic sensor intended to aid in the detection of the target.
  • the terminal vehicle 6 includes all the usual means (not described further), which are necessary to achieve this terminal flight.
  • the cover 2 is released or at least open, after separation of the different shells 3 and 4, by activating the actuating device 7, to release the terminal vehicle 6 (steering wheel) which then separates from the rest of missile 1.
  • the missile 1 is therefore provided upstream with a separable cap 2 which is intended, in particular, to thermally and mechanically protect the terminal vehicle 6.
  • This 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 terminal vehicle 6 in the terminal phase of the flight.
  • the cap 2 is mounted on the missile 1 in an operating (or protective) position.
  • the terminal vehicle 6 is mounted inside the cover 2 which is represented by dashes.
  • the shells 3 and 4 are in the process of separating, as illustrated respectively by arrows ⁇ 1 and ⁇ 2, during a phase of opening or release of the cap 2.
  • the release of the shells 3 and 4 and the impulse to generate the movements illustrated by the arrows ⁇ 1 and ⁇ 2, are generated by the actuating device 7 preferably arranged upstream of the cover 2 (inside the latter), as shown in the figures 1 and 3 .
  • This phase of opening or dropping the cover 2 allows the terminal vehicle 6 to be released.
  • the present invention can be applied more particularly to a missile 1 having a range of flight 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 missile 1 can reach several hundred degrees Celsius under the effect of the aerothermal flow, which requires providing an effective cap 2 to allow the holding and performance of structures, electronic equipment and on-board sensors.
  • the present invention can be applied to a missile 1 evolving in all cases of the flight domain (in and out of atmosphere) and for speeds ranging from subsonic to high supersonic / hypersonic.
  • the actuating device 7 allowing the shells 3 and 4 to be ejected from the missile 1 is arranged upstream of the cover 2, between the shells 3 and 4, in a plane transverse to the longitudinal axis LL of the missile 1.
  • a reference R is used associated with the pyrotechnic actuating device 7 and defined along three orthogonal axes, namely a so-called longitudinal axis X which is oriented along the actuating device 7 which is elongated, and two axes Y and Z which define a median plane XY and a transverse plane YZ.
  • the axis Z corresponds to the longitudinal axis L-L of the missile 1.
  • the front and rear adverbs are defined with respect to the direction of movement of the piston 14, which is represented by the arrow G and described below.
  • the pyrotechnic actuator 9 comprises an activatable pyrotechnic charge 12, a combustion chamber 13 arranged at the rear of the pyrotechnic actuator 9 in the same transverse plane YZ as the pyrotechnic charge 12, and a piston 14 arranged along the longitudinal axis X, the head 15 of which is in the extension of the combustion chamber 13.
  • the pyrotechnic actuator 9 is triggered by the activation of the pyrotechnic charge 12, which is carried out in the usual way, by an order given automatically by a missile control unit (not shown) 1.
  • the pyrotechnic charge 12 When the pyrotechnic charge 12 is activated, it produces an overpressure in the combustion chamber 13 which generates the displacement of the piston 14 in the direction of the arrow G.
  • the piston 14 moves until one of its ends, opposite the head 15 of the piston, called the free end 16, presses against a fixing element 17 which is fixed to the shell 3.
  • the pyrotechnic actuator 9 may, for example, be a pyrotechnic cylinder configured to contain the debris and powder residues from the pyrotechnic reaction which are liable to damage the sensor 8 of the terminal vehicle 6 or to impair its measurement capacity.
  • the pyrotechnic actuator 9 is fixed by a first end, situated at the rear of the pyrotechnic device 7, to a fixing element 18 which is fixed to the shell 4.
  • a second end of the pyrotechnic actuator 9, opposite to said first end, is free.
  • the holding rods 10A and 10B also have a first end located at the rear of the pyrotechnic device 7 and a second end situated at the front of the pyrotechnic device 7.
  • Each holding rod 10A, 10B is fixed, as specified below , by its first end to the fastening element 17 of the shell 3 and by its second end to the fastening element 18 of the shell 4.
  • the holding rods 10A and 10B provide the mechanical connection between the shells 3 and 4 of the cover 2, in particular during the flight phase of missile 1.
  • one of the two ends of each of the holding rods 10A and 10B is provided with a thread 19A, 19B which makes it possible to screw the holding rods 10A and 10B to the fixing element 17, 18 by through a nut 20A, 20B.
  • the position of the nut 20A, 20B along the thread determines the screwing of the retaining rods 10A and 10B in one of the fastening elements 17, 18 of one of the shells 3, 4, which fixes the force exerted by the hulls 3 and 4 on top of each other during the flight phase of missile 1. This force is called mechanical prestress.
  • the holding rods 10A and 10B are linked to the pyrotechnic actuator 9 by means of mechanical yokes 21A, 21B.
  • the mechanical yokes 21A and 21B are fixed on either side of the pyrotechnic actuator 9, at the level of the body of the piston 14 in the mounting position, and surround a section of retaining rods 10A and 10B.
  • the mechanical yokes 21A and 21B may correspond to lateral extensions of the pyrotechnic actuator 9.
  • each retaining rod 10A, 10B is provided with a weakening zone 22A, 22B preferably situated in the same transverse plane YZ as the free end 16 of the piston 14 in the mounting position, between the fastening element 17 and the mechanical yoke 19A, 19B.
  • Each of the weakening zones 22A and 22B corresponds to a circular recess on a longitudinal part of the holding rods 10A and 10B, which reduces their mechanical resistance.
  • a retaining element 23A, 23B for example a pin or a collar, is arranged around the retaining rod 10A, 10B, against the end of the mechanical yoke 21A, 21B closest to the weakening zone 22A, 22B.
  • This retaining element 23A, 23B retains the retaining rod 10A, 10B in the mechanical yoke 21A, 22B in the longitudinal direction X.
  • thermal insulation elements 11A, 11B, 11C, 11D are arranged on parts of the pyrotechnic actuator 9 in order to isolate it from the heat flows to which the shells 3 and 4 of the cover 2 are subjected during the flight phase .
  • a thermal insulation element 11A is located between the fastening element 18 of the shell 4 and the pyrotechnic charge 12 to prevent the heat from the shell 4 being transmitted to the pyrotechnic charge 12 and inadvertently triggering the pyrotechnic actuator 9.
  • Two other thermal insulation elements are arranged, in the form of sleeves 11B and 11C, around the sections of the holding rods 10A and 10B which pass through the mechanical yokes 21A and 21B in order to prevent the heat flows circulating between the shells 3 and 4 by means of the holding rods 10A and 10B do not pass the pyrotechnic actuator 9.
  • a thermal insulation element 11D can be arranged opposite the free end 16 of the piston 14, and attached to the fastening element 17 of the shell 3 of the missile 1.
  • the thermal insulation elements 11A, 11B, 11C, 11D protect the pyrotechnic actuator 9 by isolating only the pyrotechnic charge 12.
  • the thermal insulation elements 11A, 11B, 11C and 11D are made of one of the following materials: mica, mullite, muscovite. These materials, while being excellent thermal insulators, have sufficient hardness not to absorb the force generated by the pyrotechnic actuator 9 in order to separate the shells 3 and 4.
  • the operating mode of the actuating device is as follows.
  • the cover 2 is kept closed by means of the holding rods 10A and 10B which are fixed by their ends to fastening elements 17 and 18 of the shells 3 and 4.
  • the stability of the cap 2 depends on the mechanical preload exerted between the shells 3 and 4.
  • This mechanical preload is managed by the holding rods 10A and 10B by adjusting the position of the nut 20A, 20B along the thread of one of the ends of the holding rods 10A and 10B.
  • the cover 2 undergoes high thermal stresses during the flight phase. These heat flows circulate between the shells 3 and 4, in particular by means of the holding rods 10A and 10B which create a thermal bridge between the fastening elements 17 and 18 of the shells 3 and 4.
  • the thermal insulation elements 11A, 11B, 11C, 11D are placed judiciously between the pyrotechnic charge 12 and the fixing element 18 of the shell 4, as well as between the retaining rods 10A and 10B and the yokes mechanical 21A and 21B.
  • a signal activates the pyrotechnic charge 12 of the pyrotechnic actuator 9. There is then an overpressure in the combustion chamber 13, which generates a pushing force on the piston 14 which moves in the direction of the arrow G.
  • the piston 14 transmits the pushing force to the shell 3. Since the pyrotechnic device 7 is fixed to the two hulls 3 and 4 by means of the holding rods 10A and 10B, the hull 3 is subjected to a pushing force equal, but in opposite direction, to that acting on the hull 4.
  • the actuation device 7, as described above, is a unitary assembly, the architecture of which makes it possible, on the one hand, to fulfill the function of maintaining the stability of the cap 2, in particular during the flight phase and on the other hand the rapid ejection function of the shells 3 and 4.
  • the architecture of the actuation device 7 makes compatible the use of a pyrotechnic actuator 9 capable of generating a large force in a very short time short, despite the high temperatures to which hulls 3 and 4 are subjected.
  • the arrangement of the thermal insulation elements 11A, 11B, 11C, 11D as well as the configuration of the holding rods 10A and 10B preserve the functioning of the pyrotechnic actuator 9 by isolating it from the thermal and mechanical stresses undergone by the shells 3 and 4.
  • the cap 2 must be ejected very quickly to allow the use of the sensor 8
  • the pyrotechnic actuator 9 makes this rapid ejection possible by generating sufficient force to break the holding rods 10A and 10B, which have previously been weakened.
  • the thermal insulation elements 11A, 11B, 11C, 11D form a localized protection which makes it possible to minimize the mass and the size of the on-board actuation device 7.
  • the pyrotechnic actuating device 7 also has the advantage of being adaptable to maintaining and ejecting any removable part of the missile 1 in a high temperature environment. Finally, the actuation device 7 operates in all cases of the flight domain (in and out of the atmosphere) of a missile 1 and for speeds ranging from subsonic to high supersonic / hypersonic.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Actuator (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
EP18290030.8A 2017-04-28 2018-04-10 Dispositif d'actionnement pour l'éjection d'au moins une partie amovible de missile, en particulier d'une coiffe Active EP3396300B1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
PL18290030T PL3396300T3 (pl) 2017-04-28 2018-04-10 Urządzenie uruchamiające do odrzucania co najmniej jednej odłączalnej części pocisku, w szczególności czepca pocisku

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
FR1700467A FR3065798A1 (fr) 2017-04-28 2017-04-28 Dispositif d'actionnement pour l'ejection d'au moins une partie amovible de missile, en particulier d'une coiffe

Publications (2)

Publication Number Publication Date
EP3396300A1 EP3396300A1 (fr) 2018-10-31
EP3396300B1 true EP3396300B1 (fr) 2019-12-25

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EP18290030.8A Active EP3396300B1 (fr) 2017-04-28 2018-04-10 Dispositif d'actionnement pour l'éjection d'au moins une partie amovible de missile, en particulier d'une coiffe

Country Status (8)

Country Link
US (1) US10942015B2 (pl)
EP (1) EP3396300B1 (pl)
JP (1) JP7029470B2 (pl)
ES (1) ES2775446T3 (pl)
FR (1) FR3065798A1 (pl)
IL (1) IL269773B2 (pl)
PL (1) PL3396300T3 (pl)
WO (1) WO2018197760A1 (pl)

Families Citing this family (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112284196B (zh) * 2020-12-25 2021-04-13 星河动力(北京)空间科技有限公司 用于运载火箭的整流罩分离系统及运载火箭
CN113513951A (zh) * 2021-04-30 2021-10-19 中国工程物理研究院总体工程研究所 全包对开式头罩的连接解锁与防热系统
CN113551565B (zh) * 2021-09-18 2021-11-30 中国科学院力学研究所 一种级间段气动保形的固体火箭及分离方法
FR3138203A1 (fr) * 2022-07-21 2024-01-26 Safran Electronics & Defense Véhicule aérien à optique frontale protégée.

Family Cites Families (8)

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Publication number Priority date Publication date Assignee Title
US5235128A (en) * 1991-04-18 1993-08-10 Loral Corporation Separable missile nosecap
JP3770430B2 (ja) * 1997-06-30 2006-04-26 株式会社アイ・エイチ・アイ・エアロスペース 飛翔体のノーズフェアリング分離装置
US7082878B2 (en) 2003-07-01 2006-08-01 Raytheon Company Missile with multiple nosecones
DE102005030090B4 (de) * 2005-06-27 2007-03-22 Diehl Bgt Defence Gmbh & Co. Kg Abwerfbare Vorsatzhaube sowie Flugkörper mit abwerfbarer Vorsatzhaube
IL189089A0 (en) * 2008-01-28 2008-08-07 Rafael Advanced Defense Sys Apparatus and method for splitting and removing a shroud from an airborne vehicle
FR2947808B1 (fr) * 2009-07-09 2011-12-09 Astrium Sas Dispositif de separation lineaire douce d'une premiere piece et d'une seconde piece
FR2966919B1 (fr) * 2010-10-29 2013-11-01 Tda Armements Sas Coiffe aerodynamique secable pour munition guidee et munition guidee comportant une telle coiffe.
FR3022885B1 (fr) * 2014-06-25 2016-10-21 Mbda France Paroi structurante de missile, en particulier pour coiffe de protection thermique

Non-Patent Citations (1)

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Title
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Also Published As

Publication number Publication date
EP3396300A1 (fr) 2018-10-31
ES2775446T3 (es) 2020-07-27
US20200109929A1 (en) 2020-04-09
IL269773A (en) 2019-11-28
IL269773B1 (en) 2023-12-01
US10942015B2 (en) 2021-03-09
IL269773B2 (en) 2024-04-01
FR3065798A1 (fr) 2018-11-02
WO2018197760A1 (fr) 2018-11-01
JP7029470B2 (ja) 2022-03-03
JP2020517882A (ja) 2020-06-18
PL3396300T3 (pl) 2020-06-29

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