EP0834717A2 - Device for changing the position of the steering fins of a guided missile - Google Patents
Device for changing the position of the steering fins of a guided missile Download PDFInfo
- Publication number
- EP0834717A2 EP0834717A2 EP97116925A EP97116925A EP0834717A2 EP 0834717 A2 EP0834717 A2 EP 0834717A2 EP 97116925 A EP97116925 A EP 97116925A EP 97116925 A EP97116925 A EP 97116925A EP 0834717 A2 EP0834717 A2 EP 0834717A2
- Authority
- EP
- European Patent Office
- Prior art keywords
- rudder
- control system
- missile
- drive wheels
- sprockets
- 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
Links
- 239000000463 material Substances 0.000 claims description 2
- 230000001419 dependent effect Effects 0.000 claims 1
- 230000008878 coupling Effects 0.000 abstract description 3
- 238000010168 coupling process Methods 0.000 abstract description 3
- 238000005859 coupling reaction Methods 0.000 abstract description 3
- 230000002349 favourable effect Effects 0.000 description 1
- 230000029058 respiratory gaseous exchange Effects 0.000 description 1
- 239000007787 solid Substances 0.000 description 1
- 230000003068 static effect Effects 0.000 description 1
Images
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
Definitions
- the invention relates to a rudder control system for a guided missile the preamble of claim 1.
- Such a rudder control system is known from DE 38 27 590 C2.
- the sprockets are mounted in a rotor ring that is relative to the missile is constantly rotating around the longitudinal axis of the missile.
- the rotatable rotor ring lies between in the front part of the missile the missile tip and the engine part.
- the arrangement is also for flight stability the rudder in the front area of the missile is inappropriate.
- the invention has for its object a rudder control system of the beginning mentioned type so that it fits into the available space can be integrated at the rear of the missile.
- Fig. 1 shows a guided missile 1, which in its tail structure instead of one Fuselage constriction has a rudder control system 2. It essentially exists of four rudder blades 3 offset by 90 °, which are driven by electric motors 4 and 5, Zannkränze 7 and 8 and four drive wheels 11 can be operated. There are also two air inlets 12 and flow flaps 13 evident.
- Fig. 2 shows schematically in an exploded view the rudder control system 2 for a rudder 3.
- the sprockets 7 and 8 running on needle bearings are through Pinion 15 and 16 driven in the same direction of rotation.
- Each ring gear 7 and 8 is assigned a drive wheel 17 and 11, so that the ring gear 7 with the drive wheel 17 and the ring gear 8 cooperate with the drive wheel 11.
- the drive wheels 17 and 11 run in opposite directions Circulation directions on a rudder axis 18 freely movable.
- the rudder axis 18 with tongue and groove 21 is connected and optionally with the drive wheels 11 and 17 non-positively can be connected, whereby the rudder 3 in the desired direction deflects.
- the drive wheels 11 and 17 can be uncoupled and the rudder axis 18 can be held in the existing position by a brake 22.
- fine control can be carried out in a known manner e.g. through a change the outer geometry of the rudder 3.
- One on the rudder axis 18 attached increment encoder 23 shows the respective position of the rudder 3 with respect to its zero position defined when the missile 1 departed also serves as a signal generator for a steering control loop.
- the zero position of the rudder 3 is realized by a bolt 24, which at the launch of the missile 1 e.g. is cut by fuse wire. Changes in diameter and shape of the engine tube due to thermal influences are by rings 25 added between the missile tube and the needle bearings of the Sprockets 7 and 8 are arranged.
- the rudder control system 2 is with an aerodynamically favorable fairing 26 protected.
- FIG. 3 In the longitudinal section of FIG. 3, two of the four rudder control systems 2 are shown. From this figure, almost all of which have already been described with reference to FIG. 2 Individual parts shows, the operation of the rudder control systems 2 is particularly good to understand. It can be clearly seen that the drive wheels 11 and 17 at parallel sprockets 7 and 8 have opposite directions of rotation, whereby the rudder 3 when coupled with the clutch disks 19 adjust in the opposite direction.
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- Physics & Mathematics (AREA)
- Fluid Mechanics (AREA)
- Engineering & Computer Science (AREA)
- General Engineering & Computer Science (AREA)
- Toys (AREA)
- Retarders (AREA)
- Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
Abstract
Description
Die Erfindung betrifft ein Ruderstellsystem für einen Lenkflugkörper gemäß
dem Oberbegriff des Anspruches 1.The invention relates to a rudder control system for a guided missile
the preamble of
Ein solches Ruderstellsystem ist durch die DE 38 27 590 C2 bekannt. Dort sind die Zahnkränze in einem Rotorring gelagert, der relativ zum Flugkörper ständig eine Drehbewegung um die Längsachse des Flugkörpers ausführt. Der verdrehbare Rotorring liegt dabei im vorderen Teil des Flugkörpers zwischen der Flugkörperspitze und dem Triebwerksteil. Bei dieser Anordnung bestehen Schwierigkeiten für die Lagerung des drehbaren Rotorringes und für die statische Festigkeit des Flugkörpers. Auch ist für die Flugstabilität die Anordnung der Ruder im vorderen Bereich des Flugkörpers unzweckmäßig.Such a rudder control system is known from DE 38 27 590 C2. There the sprockets are mounted in a rotor ring that is relative to the missile is constantly rotating around the longitudinal axis of the missile. Of the rotatable rotor ring lies between in the front part of the missile the missile tip and the engine part. In this arrangement exist Difficulties for the storage of the rotatable rotor ring and for the static Missile strength. The arrangement is also for flight stability the rudder in the front area of the missile is inappropriate.
Für Flugkörper mit luftatmendem Feststofftriebwerk und zwei Lufteinlaufkanälen auf der Unterseite ist die Unterbringung der Ruderstellsysteme im Heckteil schwierig, weil durch seine Bauform bedingt am Heckteil die Einschnürung am Triebwerk nur ein geringes Maß beträgt. Es entfallt der somit üblich vorhandene Raum zwischen Marsch- und Starttriebwerk, der für die Ruderstellsysteme ausgenutzt werden könnte. Außerdem ist das Platzangebot an der vorgesehenen Rumpfstation des Trägerflugzeuges für ein Ruder sehr begrenzt, so daß an dieser Stelle ein konventionelles Ruderstellsystem nicht in Betracht kommt.For missiles with air-breathing solid engine and two air inlet channels on the bottom is the rudder positioning system in the stern difficult, because of its design, the constriction on the rear section Engine is only a small dimension. The usual existing one is therefore omitted Space between the marching and start engines, which is used for the rudder control systems could be exploited. In addition, the space is provided on the Fuselage station of the carrier aircraft for a rudder very limited, so that a conventional rudder setting system is out of the question here.
Der Erfindung liegt die Aufgabe zugrunde, ein Ruderstellsystem der eingangs genannten Art so zu gestalten, daß es in das zur Verfügung stehende Platzangebot am Heck des Flugkörpers integriert werden kann.The invention has for its object a rudder control system of the beginning mentioned type so that it fits into the available space can be integrated at the rear of the missile.
Diese Aufgabe wird durch die im Patentanspruch 1 gekennzeichenten Merkmale
gelöst. Vorteilhafte Ausgestaltungen der Erfindung sind in den Unteransprüchen
gekennzeichnet. This object is characterized by the features in
Mit der Erfindung ist es möglich, das Ruderstellsystem am hinten Ende des Flugkörpers anzuordnen, wo durch die Düsenverengung des Triebwerkes ausreichend Material vorhanden ist. Die beiden Zahnkränze mit den mit ihnen im Eingriff stehenden Antriebsrädern rotieren ständig, ohne daß der Flugkörper selbst oder ein Teil desselben rotieren muß. Mit Hilfe der zwischen den beiden in entgegengesetzter Umlaufrichtung rotierenden Antriebsrädern befindlichen Kupplung ist es möglich, jede erforderliche Ruderstellung schnell und genau einzustellen.With the invention it is possible to the rudder control system at the rear end of the Arrange missile where sufficient through the jet constriction of the engine Material is present. The two sprockets with the one in them Engaging drive wheels rotate continuously without the missile itself or part of it must rotate. With the help of between the two drive wheels rotating in the opposite direction of rotation Coupling makes it possible to quickly and accurately perform any required rudder position adjust.
Weitere Vorteile ergeben sich anhand eines nachstehend in der Zeichnung dargestellten Ausführungsbeispieles für die Erfindung. Es zeigen:
- Fig. 1
- einen Flugkörper mit einem luftatmenden Triebwerk und vier Rudern in perspektivischer Ansicht;
- Fig. 2
- den hinteren Teil des Flugkörpers nach Fig. 1, wobei für ein Ruder das Ruderstellsystem in Explosionsdarstellung gezeigt ist und
- Fig. 3
- einen Längsschnitt durch den Flugkörperteil nach Fig. 2, aus dem die Teile des Ruderstellsystems für zwei gegenüberliegende Ruder ersichtlich sind.
- Fig. 1
- a missile with an air breathing engine and four oars in a perspective view;
- Fig. 2
- the rear part of the missile according to Fig. 1, wherein the rudder control system is shown in an exploded view for a rudder and
- Fig. 3
- a longitudinal section through the missile part of FIG. 2, from which the parts of the rudder control system for two opposite rudders can be seen.
Fig. 1 zeigt einen Lenkflugkörper 1, der in seiner Heckstruktur anstelle einer
Rumpfeinschnürung ein Ruderstellsystem 2 aufweist. Es besteht im wesentlichen
aus vier um 90° versetzten Ruderblättern 3, die über Elektromotoren 4
und 5, Zannkränze 7 und 8 sowie vier Antriebsräder 11 betätigt werden können.
Weiterhin sind noch zwei Lufteinläufe 12 sowie Strömungsklappen 13
ersichtlich.Fig. 1 shows a guided
Fig. 2 zeigt schematisch in Explosionsdarstellung das Ruderstellsystem 2 für
ein Ruder 3. Die auf Nadellagern laufenden Zahnkränze 7 und 8 werden durch
Ritzel 15 und 16 in gleicher Drehrichtung angetrieben. Jedem Zahnkranz 7 und
8 ist ein Antriebsrad 17 und 11 zugeordnet, so daß also der Zahnkranz 7 mit
dem Antriebsrad 17 und der Zahnkranz 8 mit dem Antriebsrad 11 zusammenwirken.
Dadurch laufen die Antriebsräder 17 und 11 in entgegengesetzten
Umlaufrichtungen auf einer Ruderachse 18 frei beweglich. Zwischen den Antriebsrädern
11 und 17 befindet sich eine mit nicht dargestellter Ansteuerung
versehene Kupplungscheibe 19, die mit der Ruderachse 18 mit Nut und Feder
21 verbunden ist und wahlweise mit den Antriebsrädern 11 und 17 kraftschlüssig
verbunden werden kann, wodurch das Ruder 3 in gewünschter Richtung
ausschlägt. Wird über einen berechenbaren Zeitraum keine Ruderverstellung
verlangt bzw. bei hohen Geschwindigkeiten nur eine Feinlenkung benötigt,
können die Antriebsräder 11 und 17 abgekuppelt, und die Ruderachse 18
kann in der bestehenden Stellung durch eine Bremse 22 gehalten werden. In
diesem Fall kann eine Feinsteuerung in bekannter Weise z.B. durch eine Veränderung
der Außengeometrie der Ruder 3 erfolgen. Ein auf der Ruderachse
18 angebrachter Inkrementgeber 23 zeigt die jeweilige Stellung des Ruders 3
in bezug auf seine beim Abgang des Flugkörpers 1 definierte Nullage an und
dient zugleich als Signalgeber für einen Lenkregelkreis. Die Nullage des Ruders
3 wird durch einen Bolzen 24 realisiert, der beim Start des Flugkörpers 1
z.B. durch Schmelzdraht durchgetrennt wird. Durchmesser- und Formänderungen
des Triebwerkrohres durch thermische Einflüsse werden durch Ringe 25
aufgenommen, die zwischen dem Flugkörperrohr und den Nadellagern der
Zahnkränze 7 und 8 angeordnet sind. Das Ruderstellsystem 2 ist insgesamt mit
einer aerodynamisch günstigen Verkleidung 26 geschützt.Fig. 2 shows schematically in an exploded view the
In dem Längsschnitt der Fig. 3 sind zwei der vier Ruderstellsysteme 2 dargestellt.
Aus dieser Figur, die fast alle bereits anhand der Fig. 2 beschriebenen
Einzelteile zeigt, ist die Wirkungsweise der Ruderstellsysteme 2 besonders gut
zu verstehen. Es ist klar ersichtlich, daß die Antriebsräder 11 und 17 bei
gleichlaufenden Zahnkränzen 7 und 8 gegenläufige Umlaufrichtungen aufweisen,
wodurch sie beim Einkoppeln mit den Kupplungsscheiben 19 die Ruder 3
in entgegengesetzte Richtung verstellen.In the longitudinal section of FIG. 3, two of the four
Claims (6)
Applications Claiming Priority (2)
Application Number | Priority Date | Filing Date | Title |
---|---|---|---|
DE19640540A DE19640540C1 (en) | 1996-10-01 | 1996-10-01 | Rudder control system for a guided missile |
DE19640540 | 1996-10-01 |
Publications (3)
Publication Number | Publication Date |
---|---|
EP0834717A2 true EP0834717A2 (en) | 1998-04-08 |
EP0834717A3 EP0834717A3 (en) | 1999-05-06 |
EP0834717B1 EP0834717B1 (en) | 2000-12-13 |
Family
ID=7807603
Family Applications (1)
Application Number | Title | Priority Date | Filing Date |
---|---|---|---|
EP97116925A Expired - Lifetime EP0834717B1 (en) | 1996-10-01 | 1997-09-30 | Device for changing the position of the steering fins of a guided missile |
Country Status (3)
Country | Link |
---|---|
US (1) | US5975461A (en) |
EP (1) | EP0834717B1 (en) |
DE (2) | DE19640540C1 (en) |
Families Citing this family (23)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
US6073880A (en) * | 1998-05-18 | 2000-06-13 | Versatron, Inc. | Integrated missile fin deployment system |
AUPR303501A0 (en) * | 2001-02-09 | 2001-03-08 | Kusic, Tom | Spiralling missile |
US6502785B1 (en) | 1999-11-17 | 2003-01-07 | Lockheed Martin Corporation | Three axis flap control system |
US6708923B2 (en) | 2000-06-26 | 2004-03-23 | Tom Kusic | Aircraft spiralling mechanism |
DE10149087B4 (en) * | 2000-10-07 | 2008-04-03 | Bayern-Chemie Gesellschaft Für Flugchemische Antriebe Mbh | Rudder blade control for missiles |
SE521445C2 (en) * | 2001-03-20 | 2003-11-04 | Bofors Defence Ab | Methods for synchronizing the fine precipitation in a finely stabilized artillery grenade and a correspondingly designed artillery grenade |
ITMI20010648A1 (en) * | 2001-03-27 | 2002-09-27 | Finmeccanica S P A Alenia Dife | CONTROL GROUP FOR MISSILE AND / OR PROJECTILE DIRECTIONAL FLIGHTS |
AUPR583001A0 (en) | 2001-06-20 | 2001-07-12 | Kusic, Tom | Aircraft spiralling mechanism |
US7165742B2 (en) * | 2001-06-22 | 2007-01-23 | Tom Kusic | Aircraft spiralling mechanism - B |
US7093791B2 (en) * | 2001-06-22 | 2006-08-22 | Tom Kusic | Aircraft spiralling mechanism—c |
US7637453B2 (en) * | 2001-06-22 | 2009-12-29 | Tom Kusic | Aircraft spiraling mechanism with jet assistance - A |
US7635104B1 (en) | 2001-06-22 | 2009-12-22 | Tom Kusic | Aircraft spiraling mechanism with jet assistance—B |
US6637699B2 (en) | 2002-03-25 | 2003-10-28 | Lockheed Martin Corporation | Method and apparatus for controlling a trajectory of a projectile |
KR100470820B1 (en) * | 2002-04-19 | 2005-02-21 | 한국항공우주연구원 | rocket launcher |
DE102004061977B4 (en) * | 2004-12-23 | 2008-04-10 | Lfk-Lenkflugkörpersysteme Gmbh | Small Missile |
FR2895496B1 (en) * | 2005-12-26 | 2008-03-28 | Giat Ind Sa | DEVICE FOR MEASURING THE ANGULAR POSITION OF A FIN OR GOVERN OF A PROJECTILE AND METHOD OF MOUNTING SUCH A DEVICE |
US7642491B2 (en) | 2007-03-19 | 2010-01-05 | Tom Kusic | Aircraft spiraling mechanism with jet assistance—D |
US7791007B2 (en) * | 2007-06-21 | 2010-09-07 | Woodward Hrt, Inc. | Techniques for providing surface control to a guidable projectile |
US8530809B2 (en) * | 2011-08-03 | 2013-09-10 | Raytheon Company | Ring gear control actuation system for air-breathing rocket motors |
US8921749B1 (en) * | 2013-07-10 | 2014-12-30 | The United States Of America As Represented By The Secretary Of The Navy | Perpendicular drive mechanism for a missile control actuation system |
US11300390B1 (en) | 2018-03-05 | 2022-04-12 | Dynamic Structures And Materials, Llc | Control surface deployment apparatus and method of use |
KR101892900B1 (en) * | 2018-04-26 | 2018-08-28 | 국방과학연구소 | Measuring apparatus for control wing of guided missile and method thereof |
DE102018133113A1 (en) * | 2018-12-20 | 2020-06-25 | Rheinmetall Air Defence Ag | Guided missile with a plurality of wings rotatable by means of a drive arrangement with at least one actuator and at least one planetary gear |
Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE3827590A1 (en) | 1988-08-13 | 1990-02-22 | Messerschmitt Boelkow Blohm | MISSILE |
Family Cites Families (8)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE6606456U (en) * | 1963-06-14 | 1970-09-24 | Telefunken Patent | DEVICE FOR ADJUSTING THE CONTROL DEVICE OF A STEERING BODY |
US4438893A (en) * | 1973-08-10 | 1984-03-27 | Sanders Associates, Inc. | Prime power source and control for a guided projectile |
US4029270A (en) * | 1975-08-11 | 1977-06-14 | General Dynamics Corporation | Mechanical roll rate stabilizer for a rolling missile |
US4274610A (en) * | 1978-07-14 | 1981-06-23 | General Dynamics, Pomona Division | Jet tab control mechanism for thrust vector control |
JPH01300199A (en) * | 1988-05-27 | 1989-12-04 | Mitsubishi Electric Corp | Rear wing steering guided missile |
US5806791A (en) * | 1995-05-26 | 1998-09-15 | Raytheon Company | Missile jet vane control system and method |
US5829715A (en) * | 1996-04-19 | 1998-11-03 | Lockheed Martin Vought Systems Corp. | Multi-axis unfolding mechanism with rate controlled synchronized movement |
US5662290A (en) * | 1996-07-15 | 1997-09-02 | Versatron Corporation | Mechanism for thrust vector control using multiple nozzles |
-
1996
- 1996-10-01 DE DE19640540A patent/DE19640540C1/en not_active Expired - Fee Related
-
1997
- 1997-09-30 EP EP97116925A patent/EP0834717B1/en not_active Expired - Lifetime
- 1997-09-30 DE DE59702755T patent/DE59702755D1/en not_active Expired - Lifetime
- 1997-09-30 US US08/941,581 patent/US5975461A/en not_active Expired - Lifetime
Patent Citations (1)
Publication number | Priority date | Publication date | Assignee | Title |
---|---|---|---|---|
DE3827590A1 (en) | 1988-08-13 | 1990-02-22 | Messerschmitt Boelkow Blohm | MISSILE |
Also Published As
Publication number | Publication date |
---|---|
DE59702755D1 (en) | 2001-01-18 |
EP0834717A3 (en) | 1999-05-06 |
EP0834717B1 (en) | 2000-12-13 |
US5975461A (en) | 1999-11-02 |
DE19640540C1 (en) | 1998-04-02 |
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