EP2216619B1 - Dispositif d'arrêt de vol pour un corps volant - Google Patents

Dispositif d'arrêt de vol pour un corps volant Download PDF

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Publication number
EP2216619B1
EP2216619B1 EP10001071.9A EP10001071A EP2216619B1 EP 2216619 B1 EP2216619 B1 EP 2216619B1 EP 10001071 A EP10001071 A EP 10001071A EP 2216619 B1 EP2216619 B1 EP 2216619B1
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EP
European Patent Office
Prior art keywords
flight
missile
transmission
triggering
control surfaces
Prior art date
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Application number
EP10001071.9A
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German (de)
English (en)
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EP2216619A3 (fr
EP2216619A2 (fr
Inventor
Thomas Klaffert
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 Deutschland GmbH
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MBDA Deutschland GmbH
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Publication of EP2216619A3 publication Critical patent/EP2216619A3/fr
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Publication of EP2216619B1 publication Critical patent/EP2216619B1/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
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/60—Steering arrangements
    • F42B10/62—Steering by movement of flight surfaces
    • F42B10/64—Steering by movement of flight surfaces of fins
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/32—Range-reducing or range-increasing arrangements; Fall-retarding means
    • F42B10/48—Range-reducing, destabilising or braking arrangements, e.g. impact-braking arrangements; Fall-retarding means, e.g. balloons, rockets for braking or fall-retarding

Definitions

  • the present invention relates to a flight abatement device for a missile with aerodynamic control surfaces for controlling the missile according to the preamble of claim 1. Furthermore, the invention relates to a method for stopping the flight of a FugMechs with aerodynamic control surfaces.
  • this required flight abatement systems must completely or at least largely with regard to signal transmission, control and actuation can work independently. Only the power supply can be used in common, if buffering can be realized by switchable energy storage. In order to achieve a very low failure rate, it may be necessary to provide redundancy for a flight abort system. The same also applies to a telemetry system for condition monitoring of the flight cancellation system. If regular self-tests of all subsystems are required, it is also necessary to design the flight cancellation system reversibly.
  • a generic missile, in which aerodynamic control surfaces are adjustable by means of a transmission rod having a rudder drive is, for example, from DE 196 35 847 A1 known.
  • the US Pat. No. 3,352,513 forms the starting point for the preamble of claim 1 and for claim 8 and shows and describes a rotor drive of a helicopter, in which a drive shaft of the rotor drive can be severed by means of a pyrotechnic explosive device. It also describes that individual rotor blades can be separated from the rotor head by means of pyrotechnic explosive devices.
  • a pyrotechnic cutting cord can also parts of the missile, for example, rudder, wing or gear parts cut. Again, the same disadvantages apply, which have already been mentioned above in connection with the other pyrotechnic process.
  • a transverse thrust or counter thrust can be generated which deflects the missile from its predetermined path.
  • Such devices are currently mainly used in rescue capsules and landers in space technology application.
  • a disadvantage of this concept is the requirement of an additional Fuel supply, which must be structurally implemented according to the environmental requirements, for example with regard to the storage of missiles, in the missile.
  • an additional engine or the provision of an impulse charge increases the technical and design effort and the mass of the missile.
  • the additional engine must be able to be turned off and re-ignited as desired to ensure reversibility for self-tests.
  • Impulse charges must be provided in sufficient numbers on the missile to deflect the missile from its current trajectory in a preselected direction.
  • Such a method for canceling a flight is feasible for cruise missiles with turbo jet engine or liquid engine, provided that the remaining distance to the target and the area of the security areas in which the missile can be deliberately crashed, are sufficiently large.
  • This method is not suitable for missiles with solid propulsion engines, which however are used very frequently especially in military unmanned missiles. Also, this procedure of shutting down the engine is irreversible if the engine is not re-ignited.
  • This method is also used in aircraft to shorten the braking distance when landing.
  • it is only limitedly suitable for unmanned military missiles with non-disconnectable solid propulsion engine because the parachute brake generally does not withstand the high exhaust temperatures behind the missile.
  • a considerable space requirement in the missile is required for the provision of a brake parachute or parachute and the mass of the shield and trigger system increases the total mass of the missile significantly.
  • the complexity of a release system for parachute or brake parachute is due to the required high number of components in contradiction to the feasibility of a flight cancellation system with minimal risk of failure.
  • this solution is irreversible and therefore no self-test accessible.
  • Object of the present invention is therefore to provide a flight abatement device for a missile with aerodynamic control surfaces according to the preamble of claim 1, which works autonomously and reliably with minimized number of components and can be reliably used over several maintenance intervals of the missile.
  • a method for canceling the flight of a missile which deflects the missile quickly and reliably from its planned trajectory.
  • the object directed to the flight abatement device is achieved by the flight abort device having the features of patent claim 1.
  • a flight abatement device for a missile with aerodynamic control surfaces for controlling the missile wherein control surfaces of at least one drive means via at least one transmission are driven for adjustment, designed such that in the transmission at least one operable by a triggering active element is provided by at least a gear member of the transmission is formed, which is expandable in at least one of its kinematic dimensions, that the triggering device has at least one energy storage formed by an elastic element that the triggering device has at least one releasable locking device for the elastic element, that in at least one of its kinematic Dimensions expandable transmission member is biased by the elastic member in an expanded position and held by a locking device of the locking device in a compressed ready position.
  • the control surfaces can be operated independently of the drive device in the event of a flight abort and are brought into a deflecting the missile from the previous trajectory position.
  • the flight abatement device according to the invention can act on different control surfaces of the missile, for example on the wings, on rudders or on the front fuselage part of the missile intended duck wings (so-called canards).
  • canards duck wings
  • the locking device which locks the active element in the normal operating state of the missile, is unlocked upon activation of the flight abatement device, whereupon the elastic element suddenly relaxes and thereby the acted upon by the active control surfaces in a predetermined position, preferably a defined stop position brings, so that the Missile is deflected abruptly and reliably from its previous attitude.
  • a triggering device may have a plurality of independently operating locking devices.
  • the variability of the kinematic dimensions of the at least one gear member forming the active element is a length variability of the gear member.
  • the flight abatement device thus comprises a transmission element or a plurality of transmission elements, which are designed to be variably (for example variable in length) with regard to their kinematic dimensions.
  • a transmission element or a plurality of transmission elements which are designed to be variably (for example variable in length) with regard to their kinematic dimensions.
  • gear members designed as an active element they can be integrated in the transmission in parallel or in series.
  • the elastic element is formed by a spring.
  • a plurality of independently operable energy stores can be provided in a triggering device.
  • a simple and reliable construction of the energy storage device is the provision of a prestressed and preferably frictionless spring.
  • the triggering device is actuated by a control unit to move the active element from a standby position to a release position.
  • This control unit which can also be present redundantly, can be activated, for example, by the on-board computer of the missile if the on-board computer itself can make a decision for a flight abort.
  • the control unit can also be controlled via radio or telemetry from outside the missile independently of the on-board computer of the missile.
  • a reset device which returns a triggered active element from the release position to the ready position.
  • This reset device can be actuated by means of mechanical force applied externally or else be formed, for example, by a specially provided operating routine of the drive device for the control surfaces. By means of the reset device, it is possible to return an active element triggered, for example, after a self-test has been performed, back to the ready position.
  • a particularly suitable embodiment of the flight demolition device according to the invention is characterized in that the forming the active element at least one transmission member is formed by a telescopically extendable handlebar.
  • This handlebar may for example be tubular and receive the elastic element in its interior.
  • the telescopically extendable handlebar is preferably biased by the elastic element in the direction of the extended position and is held by the locking device in the collapsed standby position, wherein the locking device is part of the triggering device.
  • a plurality of independently operable locking devices can be provided.
  • the individual components of the flight abatement device according to the invention can continue to be integrated into a telemetry system for the missile to increase the reliability so that the functionality of the individual components can be monitored telemetrically from the outside.
  • a flight abort device according to the invention downstream of any existing locking system or unlocking system for the control surfaces, ie between the locking system and the unlocking system and the control surface is integrated.
  • a completely self-sufficient functioning of the flight abatement device is achieved independently of a possibly made locking the acted upon by the action mechanism control surface and independent of the set by the drive means for the control surface state.
  • At least one of the control surfaces is brought by the at least one in at least one of its kinematic dimensions expandable gear member in a deflected position to deflect the missile from its intended trajectory.
  • the at least one control surface is abruptly brought into a deflected position.
  • Fig. 1 shows a schematic representation of an inventive flight demolition device.
  • a control surface 1 is mechanically connected via a transmission 2 with a drive device 3.
  • the control surface 1 is pivotally mounted about an axis 10.
  • a pivot lever 12 is further attached, which is fixedly connected at its one end to the control surface 1 and which is pivotally connected at its other, free end to a first end of a gear member 20 of the transmission 2 via a first joint 22.
  • the rod-like gear member 20 is pivotally mounted at its second end with a second hinge 24 to a drive sleeve 30 of the drive means 3.
  • the drive sleeve 30 is provided with a through hole having an internal thread which is in threaded engagement with an external thread of a drive spindle 32 of the drive device 3.
  • the drive spindle 32 is connected to the drive shaft of a drive motor 34 of the drive device 3.
  • the drive motor 34 actuates the drive spindle 32 in a desired direction of rotation about the axis 33 of the drive spindle 32.
  • the angle of attack of the control surface 1 and thus the aerodynamic steering behavior of the control surface 1 can thus be influenced by selective actuation of the drive motor 34.
  • the gear member 20 and thus the active element 4 is designed as a telescopically extendable handlebar 40, which is designed for the transmission of tensile and compressive forces.
  • the handlebar 40 has a first, outer telescopic tube 42 and a second, inner telescopic tube 44.
  • the inner telescopic tube 44 is axially displaceably mounted by means of an annular bearing 43 in the outer telescopic tube 42, as particularly well in Fig. 2 can be seen.
  • an elastic element 5 formed by a coiled compression spring 50 is received as an energy store 5.
  • the compression spring 50 is supported at its first end on the first joint 22 receiving the head end 21 of the gear member 20 from.
  • the second end of the compression spring 50 is supported against a second end 24 receiving and connected to the outer telescopic tube 42 foot 23 of the gear member 20 from.
  • the compression spring 50 is thus anxious to expand the telescopically extendable handlebar 40, the means push out the inner telescopic tube 44 from the outer telescopic tube 42.
  • a releasable locking device 60 is provided, which is mounted in the region of the open end of the outer telescopic tube 42 on the outer telescopic tube 42 and part of a triggering device 6 for the active element 4.
  • the locking device 60 has a radially displaceable slide as a locking device 64, the radially inner end in the in Fig. 1 shown ready position of the active element 4 with an annular latching projection 62 which is provided on the outer circumference of the inner telescopic tube in the vicinity of the head end 21.
  • the annular detent projection 62 and the locking device 64 thus hold the telescopically extendable handlebar 40 in the ready position.
  • a control unit 66 is provided to actuate the triggering device 6.
  • Fig. 1 shows the flight abatement device according to the invention in the standby position
  • Fig. 2 shows Fig. 2 the flight demolition device in the release position.
  • the locking device 60 of the locking device 60 is moved from its locked with the annular locking projection 62 position on a command from the control unit 66 radially outward, as indicated by the arrow E in Fig. 2 is symbolized.
  • the radially inner end of the locking device 64 gets out of engagement with the annular latching projection 62 and thus releases the inner telescopic tube 44.
  • the inner telescopic tube 44 is abruptly moved from the outer telescopic tube 42 in the expansion position of the telescopically extendable handlebar 40.
  • This sudden expansion of the handlebar 40 causes the control surface 1 also abruptly about its axis 10 in an extreme position ( Fig. 2 ) pivots, in which the pivot lever 12 comes against a missile-fixed stop 14 to the plant.
  • This extreme deflection of the control surface 1 in turn causes that acting on the control surface 1 aerodynamic forces abruptly deflect the missile from the previous trajectory.
  • the facing ends of the locking device 64 and the annular latching projection 62 are wedge-shaped, so that the locking device when the telescopically extendable handlebar 40 from the in Fig. 2 shown expansion position is compressed again over the annular locking projection 62 and then move back in Fig. 1 shown latched position can take.
  • the drive means 3 the outer telescopic tube 42 under compression of the compression spring 50 so far move against the pivot lever 12 supported on the stop 14 inner telescopic tube 44, that this penetrates into the outer telescopic tube 42 to the locking of the locking device 64 with the locking projection 62 ,

Landscapes

  • Physics & Mathematics (AREA)
  • Fluid Mechanics (AREA)
  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)

Claims (9)

  1. Dispositif d'interruption de vol pour un missile muni de gouvernes aérodynamiques (1) servant à commander le missile,
    - les gouvernes (1) pouvant être entraînées pour leur positionnement par au moins un système d'entraînement (3) par le biais d'au moins un engrenage (2),
    caractérisé en ce
    - qu'au moins un élément actif (4) actionnable par un dispositif de déclenchement (6) est présent dans l'engrenage (2), lequel est formé par au moins un organe d'engrenage (20) de l'engrenage (2) qui peut être expansé dans au moins l'une de ses dimensions cinématiques ;
    - que le dispositif de déclenchement (6) possède au moins un accumulateur d'énergie (5) formé par un élément élastique ;
    - que le dispositif de déclenchement (6) possède au moins un dispositif de verrouillage (60) libérable pour l'élément élastique ;
    - que l'organe d'engrenage (20) qui peut être expansé dans au moins l'une de ses dimensions cinématiques est précontraint par l'élément élastique dans une position expansée et maintenu dans une position d'attente comprimée par un système d'arrêt (64) du dispositif de verrouillage (60).
  2. Dispositif d'interruption de vol selon la revendication 1, caractérisé en ce que l'au moins un organe d'engrenage (20) formant l'élément actif (4) est à longueur variable.
  3. Dispositif d'interruption de vol selon l'une des revendications précédentes, caractérisé en ce que l'élément élastique est formé par un ressort (50).
  4. Dispositif d'interruption de vol selon l'une des revendications précédentes, caractérisé en ce que le dispositif de déclenchement (6) peut être actionné par une unité de commande (66) en vue de déplacer l'élément actif (4) de la position d'attente en une position de déclenchement.
  5. Dispositif d'interruption de vol selon l'une des revendications précédentes, caractérisé en ce qu'il existe un dispositif de rappel qui ramène un élément actif (4) déclenché de la position de déclenchement en la position d'attente.
  6. Dispositif d'interruption de vol selon l'une des revendications précédentes, caractérisé en ce que l'au moins un organe d'engrenage (20) formant l'élément actif (4) est formé par une barre directrice (40) pouvant être sortie à la manière d'un télescope.
  7. Dispositif d'interruption de vol selon la revendication 6, caractérisé en ce que la barre directrice (40) pouvant être sortie à la manière d'un télescope est précontrainte par l'élément élastique dans la position sortie, et en ce que la barre directrice (40) pouvant être sortie est maintenue dans la position d'attente télescopée par le système d'arrêt (64) du dispositif de verrouillage (60).
  8. Procédé d'interruption de vol d'un missile muni de gouvernes aérodynamiques (1) au moyen d'un dispositif d'interruption de vol selon l'une des revendications précédentes, procédé selon lequel au moins l'une des gouvernes (1) est amenée dans une position déployée au moyen de l'au moins un organe d'engrenage (20) qui peut être expansé dans au moins l'une de ses dimensions cinématiques afin de dévier le missile de sa trajectoire de vol prévue.
  9. Procédé selon la revendication 8, caractérisé en ce que l'au moins une des gouvernes (1) sont amenées brusquement dans une position déployée.
EP10001071.9A 2009-02-05 2010-02-03 Dispositif d'arrêt de vol pour un corps volant Active EP2216619B1 (fr)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
DE200910007731 DE102009007731A1 (de) 2009-02-05 2009-02-05 Flugabbruchvorrichtung für einen Flugkörper

Publications (3)

Publication Number Publication Date
EP2216619A2 EP2216619A2 (fr) 2010-08-11
EP2216619A3 EP2216619A3 (fr) 2014-03-19
EP2216619B1 true EP2216619B1 (fr) 2016-12-14

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ID=41785918

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EP10001071.9A Active EP2216619B1 (fr) 2009-02-05 2010-02-03 Dispositif d'arrêt de vol pour un corps volant

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EP (1) EP2216619B1 (fr)
DE (1) DE102009007731A1 (fr)

Families Citing this family (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102020119446B4 (de) 2020-05-18 2022-01-05 Deutsches Zentrum für Luft- und Raumfahrt e.V. Risikominimierungsvorrichtung für den Flugabbruch eines Flugsystems, Verfahren zur Risikominimierung beim Flugabbruch eines Flugsystems und Flugsystem
US20240060552A1 (en) * 2022-08-17 2024-02-22 Raytheon Company Sliding pin-yoke mechanism

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3352513A (en) * 1966-03-25 1967-11-14 George W Baker Helicopter rotor shaft and rotor blade severing means
DE4135557C2 (de) * 1991-10-29 1999-05-06 Diehl Stiftung & Co Ruderstelleinrichtung
DE19635847C2 (de) * 1996-09-04 1998-07-16 Daimler Benz Aerospace Ag Lenkflugkörper mit Staustrahlantrieb

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
None *

Also Published As

Publication number Publication date
DE102009007731A1 (de) 2010-08-19
EP2216619A3 (fr) 2014-03-19
EP2216619A2 (fr) 2010-08-11

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