EP3396300A1 - Betätigungsvorrichtung für den abwurf mindestens eines lösbaren teils eines flugkörpers, insbesondere einer haube - Google Patents

Betätigungsvorrichtung für den abwurf mindestens eines lösbaren teils eines flugkörpers, insbesondere einer haube Download PDF

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Publication number
EP3396300A1
EP3396300A1 EP18290030.8A EP18290030A EP3396300A1 EP 3396300 A1 EP3396300 A1 EP 3396300A1 EP 18290030 A EP18290030 A EP 18290030A EP 3396300 A1 EP3396300 A1 EP 3396300A1
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EP
European Patent Office
Prior art keywords
missile
holding rod
piston
pyrotechnic
thermal insulation
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.)
Granted
Application number
EP18290030.8A
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English (en)
French (fr)
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EP3396300B1 (de
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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Filing date
Publication date
Application filed by MBDA France SAS filed Critical MBDA France SAS
Priority to PL18290030T priority Critical patent/PL3396300T3/pl
Publication of EP3396300A1 publication Critical patent/EP3396300A1/de
Application granted granted Critical
Publication of EP3396300B1 publication Critical patent/EP3396300B1/de
Active legal-status Critical Current
Anticipated expiration legal-status Critical

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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 actuating device for ejecting at least one removable part of a missile, and a missile provided with at least one such actuating device.
  • the present invention can be applied to a missile comprising at least one propellable propellant 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 homing device and capable of being sensitive to temperature.
  • the present invention is applicable 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 to the strength and performance of 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 ejection of the cap is implemented by an actuator configured to generate a sufficient force to separate the individual shells in a very short time to make the sensor quickly operational and to avoid any disruption of the performance of the missile during the ejection phase of the cap.
  • the actuating device must take into account the thermal and mechanical stresses to which the individual hulls are subjected before the terminal phase of flight.
  • One solution could be to use a pyrotechnic actuator such as an ejector pyrotechnic bolt, to generate the force required to separate the individual shells in very short times.
  • a pyrotechnic actuator such as an ejector pyrotechnic bolt
  • the temperatures of several hundred degrees Celsius to which the individual shells are subjected may degrade the operation of the pyrotechnic actuator fixed thereto, or even trigger it inadvertently.
  • the ejected products and the blast effect of the pyrotechnic reaction are liable to damage the sensor of the terminal vehicle or to impair its measuring capacity by deposition of powder residues for example. This solution is therefore not applicable.
  • the present invention aims to overcome these disadvantages. It relates to an actuating device for ejecting 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 holding rod.
  • a first end of said at least one holding rod and one end of said pyrotechnic actuator are intended to be fixed on a missile element and a second end, opposite said first end of said at least one holding rod, is intended to be fixed to said removable part of the missile.
  • an actuating device for ejecting a removable part of a missile, 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 holding 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 minimizes the weight and bulk of the onboard actuator.
  • the actuating device ensures mechanical support during the flight phase.
  • the pyrotechnic actuator being fixed solely to the removable part, preferably a cap shell, at one of its ends, the device actuator is provided with one or more holding rods which provide the mechanical connection between this removable portion and a fastener, for example two individual shells of a cap.
  • these holding rods are configured to withstand 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 integral part of said pyrotechnic actuator via a mechanical clevis, which ensures, for example, a better stability of the device against mechanical stresses during the flight phase of the missile and ejection of the cap.
  • said at least one holding rod has a zone of weakness, which is preferably located near the free end of the piston.
  • said at least one holding rod is provided with at least one retaining element, located at the level of the mechanical clevis.
  • This retention element is advantageously arranged to prevent any translational movement of the at least one holding rod relative to the pyrotechnic actuator.
  • said at least one holding 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 at the level of the mechanical clevis.
  • the advantageous arrangement of said at least one sleeve contributes to the thermal insulation of said pyrotechnic actuator.
  • said thermal insulation elements may be made of a material of mica, mullite or muscovite type.
  • the second end of said holding rod is advantageously provided with a thread, arranged to allow the fixing of said holding 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 missile structure and a second end, opposite the first end, to an attachment element of a removable part of the missile.
  • this removable part may 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 portion represents one of said individual shells and said second removable portion represents the other individual shell.
  • the actuating device is configured to simultaneously separate and discard the two individual shells in order to eject them from the missile.
  • At least one thermal insulation element is advantageously fixed on a fastening element of at least one of said removable parts of the missile, and arranged opposite the free end of said piston.
  • the Figures 1 and 2 schematically show an example of missile with cap, respectively, during the flight phase and during the ejection phase.
  • the figure 3 shows the arrangement of a particular embodiment of an actuating device on one of the individual shells of the cap.
  • the figures 4 and 5 are schematic views, respectively, in perspective and in median section of the actuating device.
  • the present invention applies to a missile 1 diagrammatically represented on the Figures 1 and 2 , which is provided at the front (in the direction of movement F of said missile 1) of a cap (protective) 2 having a plurality of removable parts, in this case a plurality of shells 3, 4.
  • the present invention relates to a actuating device 7 for the ejection of the cap 2.
  • the present invention can be applied to any type of missile 1 having at least one removable portion to be ejected.
  • the missile 1 with longitudinal axis LL comprises at least one propellable stage 5 releasable 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, forming for example part of a homing device and capable of being 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.
  • the propulsion stage or stages 5 of such a missile 1 are intended for the propulsion of said missile 1, from the firing to 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.
  • the terminal vehicle 6 includes all the usual means (not further described), which are necessary to achieve this terminal flight.
  • the cap 2 is released or at least open, after separation of the different shells 3 and 4, by the activation of the actuating device 7, to release the terminal vehicle 6 (steering wheel) which then separates from the rest of the missile 1.
  • the missile 1 is therefore provided upstream of a separable cap 2 which is intended, in particular, to thermally and mechanically protect the vehicle terminal 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 position (or protection).
  • the vehicle terminal 6 is mounted inside the cap 2 which is represented by dashes.
  • the shells 3 and 4 are separating, as illustrated respectively by arrows ⁇ 1 and ⁇ 2, during an opening or unloading phase of the cap 2.
  • the release of the shells 3 and 4 and the pulse to generate the movements illustrated by the arrows ⁇ 1 and ⁇ 2, are generated by the actuating device 7 preferably arranged upstream of the cap 2 (inside the latter), as shown in FIGS. figures 1 and 3 .
  • This phase of opening or releasing of the cap 2 allows the release of the terminal vehicle 6.
  • the present invention can be applied more particularly to a missile 1 having a field of flight remaining in the atmosphere and 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 cap 2 effective to allow the holding and performance of structures, electronic equipment and embedded sensors.
  • the present invention can be applied to a missile 1 evolving in all cases of the flight domain (in and out of the atmosphere) and for speeds ranging from subsonic to supersonic high / hypersonic.
  • the actuating device 7 for ejecting the hulls 3 and 4 of the missile 1 is arranged upstream of the cap 2, between the hulls 3 and 4, in a plane transverse to the longitudinal axis LL of the missile 1.
  • the Z axis 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, whose head 15 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 manner, by an order given automatically by a control unit (not shown) of the missile 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 to the piston head 15, said free end 16, bears against a fastening element 17 which is fixed to the shell 3.
  • the pyrotechnic actuator 9 may, for example, be a pyrotechnic jack configured to contain the debris and the powder residues of the pyrotechnic reaction that are likely 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, located at the rear of the pyrotechnic device 7, to a fastening 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 located 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 cap 2, especially during the flight phase of the 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 fastening element 17, 18 by via a nut 20A, 20B.
  • the position of the nut 20A, 20B along the thread determines the screwing of the holding 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 each other during the flight phase of the missile 1. This force is called mechanical prestressing.
  • the holding rods 10A and 10B are connected to the pyrotechnic actuator 9 via mechanical clevises 21A, 21B.
  • mechanical clevises 21A and 21B are fixed on either side of the pyrotechnic actuator 9, at the piston body 14 in the mounting position, and surround a section of holding rods 10A and 10B.
  • the mechanical clevises 21A and 21B may correspond to lateral extensions of the pyrotechnic actuator 9.
  • each holding rod 10A, 10B is provided with an embrittlement zone 22A, 22B located, preferably, in the same transverse plane YZ as the free end 16 of the piston 14 in the mounting position, between the fixing element 17 and the mechanical clevis 19A, 19B.
  • Each of the weakening zones 22A and 22B corresponds to a circular recess on a longitudinal portion of the holding rods 10A and 10B, which reduces their mechanical strength.
  • a retaining element 23A, 23B for example a pin or a collar, is arranged around the holding rod 10A, 10B, against the end of the mechanical clevis 21A, 21B the closest to the weakening zone 22A, 22B.
  • This retaining element 23A, 23B retains the holding rod 10A, 10B in the mechanical clevis 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 cap 2 are subjected during the flight phase. .
  • a thermal insulation element 11A is located between the fixing element 18 of the shell 4 and the pyrotechnic charge 12 to prevent the heat of the shell 4 is transmitted to the pyrotechnic charge 12 and inadvertently triggers 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 clevises 21A and 21B to prevent the heat flows circulating between them. shells 3 and 4 through 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 hull 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 heat insulators, have a hardness sufficient not to damp the force generated by the pyrotechnic actuator 9 to separate the shells 3 and 4.
  • the operating mode of the actuating device is as follows.
  • the cap 2 is kept closed by means of the holding rods 10A and 10B which are fixed at their ends to fastening elements 17 and 18 of the shells 3 and 4.
  • the stability of the cap 2 depends on the mechanical prestressing 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 end of the holding rods 10A and 10B.
  • the cap 2 undergoes high thermal stress 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 arranged judiciously between the pyrotechnic charge 12 and the fastening element 18 of the shell 4, and between the holding rods 10A and 10B and the clevises mechanical 21A and 21 B.
  • a signal activates the pyrotechnic charge 12 of the pyrotechnic actuator 9.
  • An overpressure then occurs in the combustion chamber 13, which generates a thrust force on the piston 14 which moves in the direction of the arrow G.
  • the piston 14 transmits the thrust force to the shell 3. Since the pyrotechnic device 7 is attached to the two shells 3 and 4 through the holding rods 10A and 10B, the shell 3 is subjected to an equal thrust force, but in the opposite direction, to that acting on the shell 4.
  • the retaining elements 23A and 23B arranged on the holding rods 10A and 10B at the mechanical clevises 21A and 21B, block any translational movement of the rods relative to the pyrotechnic actuator 9, the shells 3 and 4 separate and deviate from one another simultaneously by pivoting about rotating elements 24, for example hinges. This leads to the ejection of the hulls 3 and 4 of the missile 1.
  • the actuating device 7, as described above, is a unitary unit whose architecture makes it possible to fulfill, on the one hand, the function of maintaining the stability of the cap 2, in particular during the flight phase and secondly the fast ejection function of the shells 3 and 4.
  • the architecture of the actuating 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 the 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 operation of the pyrotechnic actuator 9 by isolating it from the thermal and mechanical stresses experienced by the shells 3 and 4.
  • the cap 2 must be ejected very rapidly to allow the use of the sensor 8
  • the pyrotechnic actuator 9 makes this rapid ejection possible by generating a force sufficient to break the previously weakened holding rods 10A and 10B.
  • the thermal insulation elements 11A, 11B, 11C, 11D form a localized protection which makes it possible to minimize the mass and the bulk of the on-board actuating device 7.
  • the pyrotechnic actuating device 7 also has the advantage of being adaptable to the holding and ejection of any removable part of missile 1 in a high temperature environment. Finally, the actuating device 7 operates in all cases of the flight envelope (in and out of the atmosphere) of a missile 1 and for speeds ranging from subsonic to supersonic high / 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 Betätigungsvorrichtung für den abwurf mindestens eines lösbaren teils eines flugkörpers, insbesondere einer haube Active EP3396300B1 (de)

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 true EP3396300A1 (de) 2018-10-31
EP3396300B1 EP3396300B1 (de) 2019-12-25

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EP18290030.8A Active EP3396300B1 (de) 2017-04-28 2018-04-10 Betätigungsvorrichtung für den abwurf mindestens eines lösbaren teils eines flugkörpers, insbesondere einer haube

Country Status (8)

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

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112284196A (zh) * 2020-12-25 2021-01-29 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
WO2024017767A1 (fr) * 2022-07-21 2024-01-25 Safran Electronics & Defense Véhicule aérien à optique frontale protégée

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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

Citations (3)

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EP1685362A2 (de) * 2003-11-17 2006-08-02 Raytheon Company Geschoss mit mehreren nasenkegeln
WO2009095910A2 (en) * 2008-01-28 2009-08-06 Rafael Advanced Defense Systems Ltd. Apparatus and method for splitting and removing a shroud from an airborne vehicle
EP2960619A1 (de) * 2014-06-25 2015-12-30 MBDA France Strukturelle wand eines flugkörpers, insbesondere für eine thermische schutzhaube

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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
DE102005030090B4 (de) * 2005-06-27 2007-03-22 Diehl Bgt Defence Gmbh & Co. Kg Abwerfbare Vorsatzhaube sowie Flugkörper mit abwerfbarer Vorsatzhaube
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.

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1685362A2 (de) * 2003-11-17 2006-08-02 Raytheon Company Geschoss mit mehreren nasenkegeln
WO2009095910A2 (en) * 2008-01-28 2009-08-06 Rafael Advanced Defense Systems Ltd. Apparatus and method for splitting and removing a shroud from an airborne vehicle
EP2960619A1 (de) * 2014-06-25 2015-12-30 MBDA France Strukturelle wand eines flugkörpers, insbesondere für eine thermische schutzhaube

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112284196A (zh) * 2020-12-25 2021-01-29 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
WO2024017767A1 (fr) * 2022-07-21 2024-01-25 Safran Electronics & Defense Véhicule aérien à optique frontale protégée
FR3138203A1 (fr) * 2022-07-21 2024-01-26 Safran Electronics & Defense Véhicule aérien à optique frontale protégée.

Also Published As

Publication number Publication date
US10942015B2 (en) 2021-03-09
WO2018197760A1 (fr) 2018-11-01
IL269773B2 (en) 2024-04-01
FR3065798A1 (fr) 2018-11-02
US20200109929A1 (en) 2020-04-09
EP3396300B1 (de) 2019-12-25
JP7029470B2 (ja) 2022-03-03
ES2775446T3 (es) 2020-07-27
PL3396300T3 (pl) 2020-06-29
IL269773A (en) 2019-11-28
JP2020517882A (ja) 2020-06-18
IL269773B1 (en) 2023-12-01

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