EP0238724B1 - Missile - Google Patents

Missile Download PDF

Info

Publication number
EP0238724B1
EP0238724B1 EP86117113A EP86117113A EP0238724B1 EP 0238724 B1 EP0238724 B1 EP 0238724B1 EP 86117113 A EP86117113 A EP 86117113A EP 86117113 A EP86117113 A EP 86117113A EP 0238724 B1 EP0238724 B1 EP 0238724B1
Authority
EP
European Patent Office
Prior art keywords
missile
fuselage
fuel
fins
outflow apertures
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.)
Expired - Lifetime
Application number
EP86117113A
Other languages
German (de)
French (fr)
Other versions
EP0238724A1 (en
Inventor
Berthold Schäfer
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.)
Deutsches Zentrum fuer Luft und Raumfahrt eV
Original Assignee
Deutsches Zentrum fuer Luft und Raumfahrt eV
Deutsche Forschungs und Versuchsanstalt fuer Luft und Raumfahrt eV DFVLR
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Deutsches Zentrum fuer Luft und Raumfahrt eV, Deutsche Forschungs und Versuchsanstalt fuer Luft und Raumfahrt eV DFVLR filed Critical Deutsches Zentrum fuer Luft und Raumfahrt eV
Publication of EP0238724A1 publication Critical patent/EP0238724A1/en
Application granted granted Critical
Publication of EP0238724B1 publication Critical patent/EP0238724B1/en
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

Links

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
    • F42B10/66—Steering by varying intensity or direction of thrust
    • F42B10/668—Injection of a fluid, e.g. a propellant, into the gas shear in a nozzle or in the boundary layer at the outer surface of a missile, e.g. to create a shock wave in a supersonic flow

Definitions

  • the invention relates to a missile, in particular a supersonic missile, with a fuselage, a plurality of guide fins at the rear fuselage end in the flight direction and a plurality of outflow openings arranged between the guide fins on the fuselage to influence the flight direction.
  • Such a missile is known from DE-A-33 40 037 Al.
  • This missile has a plurality of guide fins which form a ring stabilizer. Several outflow openings are arranged between these guide fins, from which gas under pressure exits in order to influence the direction of flight.
  • the escaping gas fills the space between two adjacent guide fins and the associated ring tail section, so that the air in this area is deflected and flows past the ring tail outside. As a result, a lateral force is exerted on the missile.
  • the escaping gas must be under a high pressure so that it represents an effective resistance to the air flowing past, even at high speeds. To control the missile, relatively large amounts of gas are required, which leads to weight and space problems in the missile.
  • DE-A-28 46 372 A1 discloses a projectile without guide fins or mechanism, on the fuselage of which several outflow openings are arranged, from which a gas or liquid jet emerges to influence the projectile flight direction and which can be ignited by an ignition device .
  • the outflowing gas or liquid jet forms a compression shock in the direction of flight in front of the outflow openings, with the result that that an area of high-tension air is created between the shock and the fuselage, which exerts a transverse force on the projectile.
  • the ignition of the control jet increases the effect on the air flow around the floor. At flight speeds of more than one mach, however, the flame which arises when the jet is ignited is blown out or carried by the air flow to behind the floor, so that the effects on the direction of flight are minimal or nonexistent.
  • the invention has for its object to provide a missile in which only small amounts of fuel are required to control the direction of flight.
  • the invention provides for an ignitable fuel to be emitted from the outflow openings in order to influence the direction of flight, and for the outflow openings to be arranged in areas in which recirculation of the air flowing along the fuselage forms as a result of shock waves emanating from the leading edge of the guide fin.
  • a fuel emerges from the outflow opening of the missile according to the invention, which ignites outside the fuselage. Without additional measures, the flame that arises when the escaping fuel is ignited would be due to the air traveling along the fuselage at supersonic speed immediately unstable, ie it is blown out immediately or the fuel is burned in the area behind the missile.
  • the outflow opening is located in an area on the fuselage in which the speed of the air flowing along the fuselage is reduced.
  • a so-called recirculation area arises from the fact that a shock wave emanating from the leading edge of a guide fin of the air passing the missile is reflected on an adjacent guide fin.
  • the shock wave interferes with the corner flow between fin and fuselage, which is much slower than the undisturbed flow due to the wall friction.
  • the interference of the shock wave and this corner flow creates a recirculation area near the guide fins.
  • the outflow opening is arranged in the form of a nozzle on the fuselage in an area in which the recirculation area is formed.
  • the flow velocity in the recirculation area is significantly lower than the speed of the air flowing along the fuselage outside of this area.
  • the flame that arises when the fuel emerging from the outflow opening is ignited can therefore be stabilized locally even at speeds of several Mach. This means that on the one hand the flame forms in the immediate vicinity of the fuselage and on the other hand that the flame is not destroyed by the air flowing past the fuselage or carried to the rear of the missile.
  • the compression shocks shock waves
  • shock waves are used for flame stabilization.
  • the fuel ignited in the recirculation area causes an increase in volume in the immediate vicinity of the fuselage.
  • This increase in volume causes a local pressure increase on the fuselage, whereby a change in the flight direction of the missile is achieved.
  • the missile is thus controlled by an increase in pressure in an area on the fuselage's fuselage that is delimited by the guide fins. Since the combustion of the fuel in the air flowing past the fuselage (external combustion) results in a large increase in volume and a high pressure increase, only small amounts of fuel are required on board the missile. As a result, the missiles can be made smaller and lighter.
  • the control of the missile can be carried out within a very short time, so that there are short reaction times.
  • the control can be used during the entire flight phase, i.e. both during the launch and the marching phase of the missile, but it is particularly effective in the supersonic range. Movable and therefore fault-prone parts, e.g. Oars are not required.
  • leading edges of the guide fins are sharpened at an acute angle on both sides. Due to this special design of the leading edges of the guide fins, these do not represent any significant air resistance for the air flowing along the fuselage. Shock waves already form at a leading edge tapering at an angle of approximately 20 ° whose strength is sufficient to generate a sufficiently trained recirculation area.
  • an ignition device is located between two adjacent guide fins, with the aid of which the fuel sprayed out of the nozzle can be ignited.
  • the ignition device is used to ignite the fuel depending on the type of fuel and the speed of the missile. If, for example, fuel is used which ignites itself at correspondingly high flight speeds of the missile (e.g. at four times the speed of sound) due to the high storage temperatures, the ignition device is only required during the launch phase of the missile. During the flight phase of the missile, the ignition device for igniting the fuel is generally not required, which simplifies the control process of the missile.
  • Another advantageous embodiment of the invention is characterized in that several outflow openings are arranged in a row between adjacent guide fins, the outer outflow openings of the row being arranged in the immediate vicinity of a guide fin.
  • the recirculation areas develop particularly in the immediate vicinity of a guide fin, since the corner flow is slowed down the most due to the friction of the air flowing along the fuselage and on the guide fin.
  • the fuel injected via the outer outflow openings of the row into these particularly well-defined recirculation areas forms a local one when it is burned stable flame. From there, the flame spreads rapidly across the entire row of outflow openings. This creates a wide flame area between the guide fins, which enables particularly effective control to be achieved.
  • the hydrogen emerging from the outflow openings self-ignites at supersonic speeds of the missile in the range of approximately 4 Mach. Due to the high damming temperatures of the air passing along the supersonic fuselage, temperatures of approx. 800 ° C are reached which lead to self-ignition of the hydrogen. In these speed ranges of the missile, the ignition device for igniting the hydrogen is not required, which simplifies the operations required to control the missile.
  • the missile according to the invention is provided with several rigid guide fins in its tail area. Nozzles are attached between adjacent guide fins, through which fuel flows into the supersonic flow of those flowing along the fuselage's fuselage Air can be injected.
  • the supersonic flow generates shock waves (compression shocks) emanating from the leading edges of the guide fins, which are reflected on the adjacent guide fin. This reflection causes interference between the shock wave and the air traveling along the fuselage of the missile, which air has a reduced speed due to its friction on the fuselage.
  • the shock interference is particularly strong in the corners formed by the fuselage and the fins. Due to the impact interference, recirculation areas form in the supersonic flow, in which a locally stabilized flame is produced when a fuel injected into these areas is burned.
  • the fuel When sprayed, the fuel only needs to have a slightly higher pressure than the air flowing along the fuselage.
  • the external combustion of the fuel in the air flow surrounding the missile leads to an increase in volume of the fuel / air mixture in the immediate vicinity of the fuselage, which results in an increase in pressure in this area.
  • This increase in pressure in the area delimited by the guide fins affects the fuselage of the missile and is thus used to control the missile.
  • the control is carried out by external combustion of a fuel.
  • This type of control of a missile is very responsive and can be used during the entire flight phase, i.e. during the launch and marching phase of the missile. Only relatively small amounts of fuel are required, as a result of which the missile can be made small in its dimensions and has a low weight.
  • the mechanism for controlling the missile has no moving parts, which makes it very reliable.
  • the missile 10 has four fins 14, 16, 18 and 20 on its fuselage 12, which are arranged at the rear end of the fuselage 12 in the direction of flight A.
  • the leading edge of a guide fin (in the figures with the reference symbol of the relevant guide fin supplemented by an F) is sharpened on both sides and tapers towards the front.
  • the radially outward-pointing side edges of the guide fins (denoted in the figures with the reference symbol of the relevant guide fin supplemented by an S) also taper to the outside.
  • a plurality of nozzles 22 are located on the fuselage 12 between the adjacent guide fins 14 and 16, with five nozzles 22 each being arranged one behind the other in a row running transversely to the direction of flight A of the missile 10 and three such rows 24, 26 and 28.
  • Nozzles 22 arranged in this way are located between all the adjacent fins of the missile 10. All of the nozzles 22 arranged in a row are located on a common circumferential circle on the fuselage 12.
  • the nozzles 22 arranged on a circumferential circle on the fuselage 12 are guided by the fins 16, 14, 20 and 18 in four groups of each divided into five nozzles. Such a group of nozzles 22 is the I., II., III. and IV. Quadrants (Fig. 2) assigned.
  • the division of the nozzles 22 in the rows 26 and 28 is corresponding.
  • Fuel is injected into the air flowing along the fuselage 12 via the nozzles 22.
  • All of the nozzles 22 arranged on the fuselage 12 are connected to a tank for fuel (likewise not shown) via lines (not shown).
  • a tank for fuel likewise not shown
  • lines not shown
  • either all the nozzles 22 of one quadrant or else the nozzles of several quadrants can be supplied with fuel.
  • the supply of the nozzles 22 with fuel is therefore selected according to quadrants.
  • an ignition device 30 Between the middle row 26 and the last row 28 (viewed in the direction of flight A) there is an ignition device 30 - for example in the form of a spark plug - for igniting the fuel emerging from the nozzles 22 of the quadrant in question.
  • the creation of a recirculation area between two adjacent guide fins is to be illustrated below using the first quadrant (using the area between guide fins 14 and 16) as an example.
  • two shock wave fronts originate from the leading edge 14F of the guide fin 14 and from the leading edge 16F of the guide fin 16.
  • One of the two extends from the leading edge 14F of the guide fin 14 outgoing shock wave fronts in the direction of the guide fin 16, while one of the two shock wave fronts starting from the front edge 16F of the guide fin 16 runs in the direction of the guide fin 14.
  • the shock waves of each shock wave front are reflected on the corresponding guide fins.
  • the shock waves interfere with the air passing along the fuselage l2 of the missile 10.
  • the air flowing along the fuselage 12 is braked due to the friction on the fuselage 12, as a result of which recirculation areas are formed in the event of interference with the shock waves.
  • the most slowed down flow of the air flowing along the fuselage occurs in the corner (corner flow) between the guide fins 14 and 16 and the fuselage 12 of the missile 10. Therefore, the most highly developed recirculation areas are also in the vicinity a guide fin.
  • fuel is injected via the nozzles 22 of one or more quadrants as required injected into the air flowing along the fuselage 12.
  • the emerging fuel is ignited with the aid of the ignition device 30, a locally stable flame being formed.
  • the combustion of the fuel causes an increase in the volume of the mixture of burned fuel located on the fuselage 12 between the guide fins 14 and 16 and air flowing along the fuselage 12.
  • This increase in volume results in an increase in pressure in the area precisely delimited by the guide fins 14 and 16.
  • the increased pressure in this area acts on the fuselage 12, as a result of which a transverse force which is directed transversely to the flight direction A of the missile 10 is generated.
  • the strength of the transverse force can be regulated via the amount of fuel exiting through the nozzles 22 of a quadrant per unit of time.
  • the pressure of the gaseous or liquid fuel emerging from the nozzles 22 is only so great that it is sufficient to allow the fuel to emerge from the fuselage 12. This pressure alone does not give the missile any significant control impulse.
  • hydrogen will self-ignite from a certain velocity of the missile 10 due to the high accumulation temperature of the air passing along the fuselage 12.
  • the ignition temperature for hydrogen is around 800 ° C. If the missile 10 has a speed greater than approx. 4 Mach under ground conditions, the temperature of the air on the missile 12 has risen to values greater than 800 ° C. due to the high accumulation temperatures, so that the hydrogen ignites reliably. At these speed ranges of the missile, those running in it can Processes during the control are simplified in such a way that the control of the corresponding ignition device need not take place with each control maneuver.

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)

Description

Die Erfindung betrifft einen Flugkörper, insbesondere einen Überschallflugkörper, mit einem Rumpf, mehreren Leitflossen am in Flugrichtung hinteren Rumpfende und mehreren zwischen den Leitflossen am Rumpf angeordneten Ausströmöffnungen zur Beeinflussung der Flugrichtung.The invention relates to a missile, in particular a supersonic missile, with a fuselage, a plurality of guide fins at the rear fuselage end in the flight direction and a plurality of outflow openings arranged between the guide fins on the fuselage to influence the flight direction.

Ein derartiger Flugkörper ist aus DE-A-33 40 037 Al bekannt. Dieser Flugkörper weist mehrere ein Ringleitwerk bildende Leitflossen auf. Zwischen diesen Leitflossen sind mehrere Ausströmöffnungen angeordnet, aus denen zur Beeinflussung der Flugrichtung unter Druck stehendes Gas austritt. Das austretende Gas füllt den Raum zwischen zwei benachbarten Leitflossen und dem zugehörigen Ringleitwerkabschnitt aus, so daß die Luft in diesem Bereich abgelenkt wird und außen am Ringleitwerk vorbeiströmt. Dadurch wird auf den Flugkörper eine Querkraft ausgeübt. Das austretende Gas muß unter einem hohen Druck stehen, damit es für die vorbeiströmende Luft auch bei hohen Geschwindigkeiten einen wirkungsvollen Widerstand darstellt. Zur Steuerung des Flugkörpers sind relativ große Gasmengen erforderlich, was zu Gewichts- und Platzproblemen im Flugkörper führt.Such a missile is known from DE-A-33 40 037 Al. This missile has a plurality of guide fins which form a ring stabilizer. Several outflow openings are arranged between these guide fins, from which gas under pressure exits in order to influence the direction of flight. The escaping gas fills the space between two adjacent guide fins and the associated ring tail section, so that the air in this area is deflected and flows past the ring tail outside. As a result, a lateral force is exerted on the missile. The escaping gas must be under a high pressure so that it represents an effective resistance to the air flowing past, even at high speeds. To control the missile, relatively large amounts of gas are required, which leads to weight and space problems in the missile.

Aus DE-A-28 46 372 Al ist ein Geschoß ohne Leitflossen bzw. -werk bekannt, an dessen Rumpf mehrere Ausströmöffnungen angeordnet sind, aus denen zur Beeinflussung der Geschoß-Flugrichtung ein Gas- oder Flüssigkeitsstrahl austritt, der durch eine Zündvorrichtung gezündet werden kann. Durch den ausströmenden Gas- oder Flüssigkeitsstrahl bildet sich in Flugrichtung vor den Ausströmöffnungen ein Verdichtungsstoß mit der Folge, daß zwischen dem Verdichtungsstoß und dem Geschoßrumpf ein Gebiet hochgespannter Luft entsteht, die auf das Geschoß eine Querkraft ausübt. Durch die Zündung des Steuerstrahls wird die Wirkung auf die Luftströmung um das Geschoß herum erhöht. Bei Fluggeschwindigkeiten von mehreren Mach wird die bei Zündung des Strahls entstehende Flamme jedoch ausgeblasen oder von dem Luftstrom bis hinter das Geschoß getragen, so daß die Auswirkungen auf die Flugrichtung nur gering oder gar nicht vorhanden sind.DE-A-28 46 372 A1 discloses a projectile without guide fins or mechanism, on the fuselage of which several outflow openings are arranged, from which a gas or liquid jet emerges to influence the projectile flight direction and which can be ignited by an ignition device . The outflowing gas or liquid jet forms a compression shock in the direction of flight in front of the outflow openings, with the result that that an area of high-tension air is created between the shock and the fuselage, which exerts a transverse force on the projectile. The ignition of the control jet increases the effect on the air flow around the floor. At flight speeds of more than one mach, however, the flame which arises when the jet is ignited is blown out or carried by the air flow to behind the floor, so that the effects on the direction of flight are minimal or nonexistent.

Der Erfindung liegt die Aufgabe zugrunde, einen Flugkörper zu schaffen, bei dem zur Steuerung der Flugrichtung nur geringe Mengen an Brennstoff notwendig sind.The invention has for its object to provide a missile in which only small amounts of fuel are required to control the direction of flight.

Zur Lösung dieser Aufgabe ist nach der Erfindung vorgesehen, zur Beeinflussung der Flugrichtung aus den Ausströmöffnungen einen zündbaren Brennstoff auszustoßen und die Ausströmöffnungen in Bereichen anzuordnen, in denen sich infolge von von den Leitflossenvorderkanten ausgehenden Stoßwellen Rezirkulationen der am Rumpf entlangstreichenden Luft bilden.In order to achieve this object, the invention provides for an ignitable fuel to be emitted from the outflow openings in order to influence the direction of flight, and for the outflow openings to be arranged in areas in which recirculation of the air flowing along the fuselage forms as a result of shock waves emanating from the leading edge of the guide fin.

Zur Steuerung der Flugrichtung tritt aus der Ausströmöffnung des erfindungsgemäßen Flugkörpers ein Brennstoff aus, der sich außerhalb des Rumpfes entzündet. Ohne zusätzliche Maßnahmen würde die bei Zündung des austretenden Brennstoffs entstehende Flamme aufgrund der mit Überschallgeschwindigkeit am Rumpf entlangstreichenden Luft sofort instabil, d.h., sie wird sofort ausgeblasen oder aber die Verbrennung des Brennstoffs erfolgt im Bereich hinter dem Flugkörper. Zur Stabilisierung der bei der Verbrennung des Brennstoffs entstehenden Flamme befindet sich die Ausströmöffnung in einem Bereich am Rumpf, in dem die Geschwindigkeit der am Rumpf entlangstreichenden Luft verringert ist. Ein derartiger Bereich, ein sogenanntes Rezirkulationsgebiet, entsteht dadurch, daß eine von der Vorderkante einer Leitflosse ausgehende Stoßwelle der am Flugkörper vorbeistreichenden Luft an einer benachbarten Leitflosse reflektiert wird. Beim Auftreffen auf die benachbarte Leitflosse interferiert die Stoßwelle mit der Eckenströmung zwischen Flosse und Rumpf, die aufgrund der Wandreibung wesentlich langsamer ist als die ungestörte Strömung. Die Interferenz der Stoßwelle und dieser Eckenströmung erzeugt ein Rezirkulationsgebiet in der Nähe der Leitflossen.To control the flight direction, a fuel emerges from the outflow opening of the missile according to the invention, which ignites outside the fuselage. Without additional measures, the flame that arises when the escaping fuel is ignited would be due to the air traveling along the fuselage at supersonic speed immediately unstable, ie it is blown out immediately or the fuel is burned in the area behind the missile. In order to stabilize the flame arising during the combustion of the fuel, the outflow opening is located in an area on the fuselage in which the speed of the air flowing along the fuselage is reduced. Such an area, a so-called recirculation area, arises from the fact that a shock wave emanating from the leading edge of a guide fin of the air passing the missile is reflected on an adjacent guide fin. When it hits the adjacent guide fin, the shock wave interferes with the corner flow between fin and fuselage, which is much slower than the undisturbed flow due to the wall friction. The interference of the shock wave and this corner flow creates a recirculation area near the guide fins.

Die Ausströmöffnung ist in Form einer Düse am Rumpf in einem Bereich angeordnet, in dem sich das Rezirkulationsgebiet ausbildet. Die Strömungsgeschwindigkeit im Rezirkulationsgebiet ist wesentlich geringer als die Geschwindigkeit der am Rumpf entlangstreichenden Luft außerhalb dieses Gebietes. Die Flamme, die bei Zündung des aus der Ausströmöffnung austretenden Brennstoffes entsteht, kann daher selbst bei Geschwindigkeiten von mehreren Mach lokal stabilisiert werden. Das bedeutet, daß sich einerseits die Flamme in unmittelbarer Nähe des Rumpfes bildet und daß andererseits die Flamme nicht von der am Rumpf vorbeiströmenden Luft zerstört bzw. bis hinter den Flugkörper getragen wird. Bei dem erfindungsgemäßen Flugkörper werden die Verdichtungsstöße (Stoßwellen) zur Flammenstabilisation verwendet.The outflow opening is arranged in the form of a nozzle on the fuselage in an area in which the recirculation area is formed. The flow velocity in the recirculation area is significantly lower than the speed of the air flowing along the fuselage outside of this area. The flame that arises when the fuel emerging from the outflow opening is ignited can therefore be stabilized locally even at speeds of several Mach. This means that on the one hand the flame forms in the immediate vicinity of the fuselage and on the other hand that the flame is not destroyed by the air flowing past the fuselage or carried to the rear of the missile. In the missile according to the invention, the compression shocks (shock waves) are used for flame stabilization.

Der im Rezirkulationsgebiet gezündetete Brennstoff bewirkt eine Volumenvergrößerung in unmittelbarer Nähe des Rumpfes. Diese Volumenvergrößerung bewirkt einen lokalen Druckanstieg am Rumpf, wodurch eine Veränderung der Flugrichtung des Flugkörpers erzielt wird. Die Steuerung des Flugkörpers erfolgt also durch Druckanstieg in einem durch die Leitflossen örtlich genau begrenzten Gebiet am Rumpf des Flugkörpers. Da sich bei der Verbrennung des Brennstoffs in der am Rumpf vorbeiströmenden Luft (Außenverbrennung) eine starke Volumenvergrößerung und ein hoher Druckanstieg ergibt, sind nur geringe Brennstoffmengen an Bord des Flugkörpers erforderlich. Dadurch können die Flugkörper kleiner und leichter ausgebildet sein. Die Steuerung des Flugkörpers kann innerhalb kürzester Zeit vorgenommen werden, so daß sich kurze Reaktionszeiten ergeben. Die Steuerung kann während der gesamten Flugphase, also sowohl während der Start- als auch der Marschphase des Flugkörpers, eingesetzt werden, besonders wirkungsvoll ist sie aber vor allem im Überschallbereich. Bewegliche und daher störungsanfällige Teile, wie z.B. Ruder, sind nicht erforderlich.The fuel ignited in the recirculation area causes an increase in volume in the immediate vicinity of the fuselage. This increase in volume causes a local pressure increase on the fuselage, whereby a change in the flight direction of the missile is achieved. The missile is thus controlled by an increase in pressure in an area on the fuselage's fuselage that is delimited by the guide fins. Since the combustion of the fuel in the air flowing past the fuselage (external combustion) results in a large increase in volume and a high pressure increase, only small amounts of fuel are required on board the missile. As a result, the missiles can be made smaller and lighter. The control of the missile can be carried out within a very short time, so that there are short reaction times. The control can be used during the entire flight phase, i.e. both during the launch and the marching phase of the missile, but it is particularly effective in the supersonic range. Movable and therefore fault-prone parts, e.g. Oars are not required.

Gemäß einer vorteilhaften Ausführungsform der Erfindung ist vorgesehen, daß die Vorderkanten der Leitflossen zu beiden Seiten spitzwinklig angeschärft sind. Durch diese besondere Ausbildung der Vorderkanten der Leitflossen stellen diese keinen wesentlichen Luftwiderstand für die am Rumpf entlangstreichende Luft dar. Bereits bei einer mit einem Winkel von ca. 20° zulaufenden Vorderkante bilden sich an dieser Stoßwellen aus, deren Stärke zur Erzeugung eines hinreichend ausgebildeten Rezirkulationsgebietes ausreichen.According to an advantageous embodiment of the invention, it is provided that the leading edges of the guide fins are sharpened at an acute angle on both sides. Due to this special design of the leading edges of the guide fins, these do not represent any significant air resistance for the air flowing along the fuselage. Shock waves already form at a leading edge tapering at an angle of approximately 20 ° whose strength is sufficient to generate a sufficiently trained recirculation area.

Gemäß einer weiteren Ausgestaltung der Erfindung befindet sich zwischen zwei benachbarten Leitflossen eine Zündvorrichtung, mit deren Hilfe der aus der Düse ausgespritzte Brennstoff entzündet werden kann. Die Zündvorrichtung wird je nach Art des Brennstoffs sowie der Geschwindigkeit des Flugkörpers zum Zünden des Brennstoffs verwendet. Wenn beispielsweise Brennstoff verwendet wird, der sich bei entsprechend hohen Fluggeschwindigkeiten des Flugkörpers (z.B. bei vierfacher Schallgeschwindigkeit) infolge der hohen Stautemperaturen selbst entzündet, wird die Zündvorrichtung nur während der Startphase des Flugkörpers benötigt. Während der Marschphase des Flugkörpers ist die Zündvorrichtung zum Entzünden des Brennstoffs in der Regel nicht erforderlich, wodurch der Steuerungsprozeß des Flugkörpers vereinfacht wird.According to a further embodiment of the invention, an ignition device is located between two adjacent guide fins, with the aid of which the fuel sprayed out of the nozzle can be ignited. The ignition device is used to ignite the fuel depending on the type of fuel and the speed of the missile. If, for example, fuel is used which ignites itself at correspondingly high flight speeds of the missile (e.g. at four times the speed of sound) due to the high storage temperatures, the ignition device is only required during the launch phase of the missile. During the flight phase of the missile, the ignition device for igniting the fuel is generally not required, which simplifies the control process of the missile.

Eine andere vorteilhafte Ausgestaltung der Erfindung ist dadurch gekennzeichnet, daß zwischen benachbarten Leitflosse jeweils mehrere Ausströmöffnungen in einer Reihe angeordnet sind, wobei die äußeren Ausströmöffnungen der Reihe jeweils in unmittelbarer Nähe einer Leitflosse angeordnet sind. Die Rezirkulationsgebiete bilden sich insbesondere in unmittelbarer Nähe einer Leitflosse aus, da die Eckenströmung aufgrund der Reibung der entlangstreichenden Luft am Rumpf und an der Leitflosse am stärksten abgebremst ist. Der über die äußeren Ausströmöffnungen der Reihe in diese besonders gut ausgeprägten Rezirkulationsgebiete eingespritzte Brennstoff bildet bei seiner Verbrennung eine lokal stabile Flamme. Von dort aus breitet sich die Flamme rasch über die gesamte Reihe der Ausströmöffnungen hinweg aus. So entsteht ein breiter Flammenbereich zwischen den Leitflossen, wodurch eine besonders wirkungsvolle Steuerung erzielt werden kann.Another advantageous embodiment of the invention is characterized in that several outflow openings are arranged in a row between adjacent guide fins, the outer outflow openings of the row being arranged in the immediate vicinity of a guide fin. The recirculation areas develop particularly in the immediate vicinity of a guide fin, since the corner flow is slowed down the most due to the friction of the air flowing along the fuselage and on the guide fin. The fuel injected via the outer outflow openings of the row into these particularly well-defined recirculation areas forms a local one when it is burned stable flame. From there, the flame spreads rapidly across the entire row of outflow openings. This creates a wide flame area between the guide fins, which enables particularly effective control to be achieved.

Gemäß einer bevorzugten Ausführungsform der Erfindung ist vorgesehen, daß mehrere hintereinanderliegende Reihen von Ausströmöffnungen vorgesehen sind. Dadurch entsteht ein besonders großflächiger Bereich zwischen zwei benachbarten Leitflossen, in dem der austretende Brennstoff verbrennt. Auch dies ermöglicht eine besonders wirkungsvolle Steuerung des Flugkörpers.According to a preferred embodiment of the invention it is provided that several rows of outflow openings are provided one behind the other. This creates a particularly large area between two adjacent guide fins in which the escaping fuel burns. This also enables a particularly effective control of the missile.

Wird, wie es bei einer weiteren Ausführungsform der Erfindung vorgesehen ist, als Brennstoff Wasserstoff verwendet, so erfolgt bei Überschallgeschwindigkeiten des Flugkörpers im Bereich von ca. 4 Mach eine Selbstentzündung des aus den Ausströmöffnungen austretenden Wasserstoffs. Aufgrund der hohen Stautemperaturen der mit Überschall am Rumpf des Flugkörpers entlangstreichenden Luft werden hierbei Temperaturen von ca. 800°C erreicht, die zur Selbstentzündung des Wasserstoffs führen. In diesen Geschwindigkeitsbereichen des Flugkörpers wird die Zündvorrichtung zum Entzünden des Wasserstoffs nicht benötigt, wodurch die zur Steuerung des Flugkörpers notwendigen Operationen vereinfacht werden.If, as is provided in a further embodiment of the invention, hydrogen is used as the fuel, the hydrogen emerging from the outflow openings self-ignites at supersonic speeds of the missile in the range of approximately 4 Mach. Due to the high damming temperatures of the air passing along the supersonic fuselage, temperatures of approx. 800 ° C are reached which lead to self-ignition of the hydrogen. In these speed ranges of the missile, the ignition device for igniting the hydrogen is not required, which simplifies the operations required to control the missile.

Der erfindungsgemäße Flugkörper wird in seinem Heckbereich mit mehreren starren Leitflossen versehen. Zwischen benachbarten Leitflossen werden Düsen angebracht, durch die Brennstoff in die Überschallströmung der am Rumpf des Flugkörpers entlangstreichenden Luft eingespritzt werden kann. Die Überschallströmung erzeugt von den Vorderkanten der Leitflossen ausgehende Stoßwellen (Verdichtungsstöße), die an der jeweils benachbarten Leitflosse reflektiert werden. Bei dieser Reflektion erfolgt eine Interferenz zwischen der Stoßwelle und der am Rumpf des Flugkörpers entlangstreichenden Luft, die aufgrund ihrer Reibung am Rumpf eine verringerte Geschwindigkeit aufweist. Die Stoßinterferenz ist besonders stark in den von dem Rumpf und den Leitflossen gebildeten Ecken. Aufgrund der Stoßinterferenz bilden sich in der Überschallströmung Rezirkulationsgebiete, in denen bei Verbrennung eines in diese Gebiete eingespritzten Brennstoffs eine lokal stabilisierte Flamme entsteht. Der Brennstoff braucht beim Ausspritzen nur einen geringfügig größeren Druck als die am Rumpf entlangströmende Luft aufzuweisen. Die Außenverbrennung des Brennstoffs in der den Flugkörper umgebenden Luftströmung führt zu einer Volumenvergrößerung des Brennstoff/Luftgemisches in unmittelbarer Nähe des Rumpfes, was in diesem Bereich einen Druckanstieg zur Folge hat. Dieser Druckanstieg in dem durch die Leitflossen örtlich genau begrenzten Gebiet wirkt sich auf den Rumpf des Flugkörpers aus und wird so zur Steuerung des Flugkörpers verwendet. Die Steuerung erfolgt also durch Außenverbrennung eines Brennstoffs. Diese Art der Steuerung eines Flugkörpers erfolgt sehr reaktionsschnell und kann während der gesamten Flugphase, also während der Start- und Marschphase des Flugkörpers eingesetzt werden. Es sind nur relativ geringe Brennstoffmengen erforderlich, wodurch der Flugkörper in seinen Abmessungen klein gestaltet werden kann und ein geringes Gewicht aufweist. Der Mechnismus zur Steuerung des Flugkörpers weist keine beweglichen Teile auf, wodurch er sehr zuverlässig arbeitet.The missile according to the invention is provided with several rigid guide fins in its tail area. Nozzles are attached between adjacent guide fins, through which fuel flows into the supersonic flow of those flowing along the fuselage's fuselage Air can be injected. The supersonic flow generates shock waves (compression shocks) emanating from the leading edges of the guide fins, which are reflected on the adjacent guide fin. This reflection causes interference between the shock wave and the air traveling along the fuselage of the missile, which air has a reduced speed due to its friction on the fuselage. The shock interference is particularly strong in the corners formed by the fuselage and the fins. Due to the impact interference, recirculation areas form in the supersonic flow, in which a locally stabilized flame is produced when a fuel injected into these areas is burned. When sprayed, the fuel only needs to have a slightly higher pressure than the air flowing along the fuselage. The external combustion of the fuel in the air flow surrounding the missile leads to an increase in volume of the fuel / air mixture in the immediate vicinity of the fuselage, which results in an increase in pressure in this area. This increase in pressure in the area delimited by the guide fins affects the fuselage of the missile and is thus used to control the missile. The control is carried out by external combustion of a fuel. This type of control of a missile is very responsive and can be used during the entire flight phase, i.e. during the launch and marching phase of the missile. Only relatively small amounts of fuel are required, as a result of which the missile can be made small in its dimensions and has a low weight. The mechanism for controlling the missile has no moving parts, which makes it very reliable.

Nachfolgend wird unter Bezugnahme auf die Figuren ein Ausführungsbeispiel der Erfindung näher erläutert.An exemplary embodiment of the invention is explained in more detail below with reference to the figures.

Es zeigen:

Fig. 1
eine Seitenansicht des Flugkörpers und
Fig. 2
einen Querschnitt entlang der Linie II-II in Fig. 1.
Show it:
Fig. 1
a side view of the missile and
Fig. 2
a cross section along the line II-II in Fig. 1st

Der Flugkörper 10 weist an seinem Rumpf 12 vier Leitflossen 14,16,18 und 20 auf, die an dem in Flugrichtung A hinteren Ende des Rumpfes 12 angeordnet sind. Die Vorderkante einer Leitflosse (in den Figuren mit dem durch ein F ergänztes Bezugszeichen der betreffenden Leitflosse bezeichnet) sind zu beiden Seiten angeschärft und laufen nach vorne hin spitz zu. Auch die radial nach außen weisenden Seitenkanten der Leitflossen (in den Figuren mit dem durch ein S ergänztes Bezugszeichen der betreffenden Leitflosse bezeichnet) laufen nach außen hin spitz zu.The missile 10 has four fins 14, 16, 18 and 20 on its fuselage 12, which are arranged at the rear end of the fuselage 12 in the direction of flight A. The leading edge of a guide fin (in the figures with the reference symbol of the relevant guide fin supplemented by an F) is sharpened on both sides and tapers towards the front. The radially outward-pointing side edges of the guide fins (denoted in the figures with the reference symbol of the relevant guide fin supplemented by an S) also taper to the outside.

Zwischen den benachbarten Leitflossen 14 und 16 befinden sich am Rumpf 12 mehrere Düsen 22, wobei jeweils fünf Düsen 22 in einer quer zur Flugrichtung A des Flugkörpers 10 verlaufenden Reihe und drei solcher Reihen 24,26 und 28 hintereinanderliegend angeordnet sind. Derart angeordnete Düsen 22 befinden sich zwischen sämtlichen zueinander benachbarten Leitflossen des Flugkörpers 10. Alle in einer Reihe angeordneten Düsen 22 befinden sich auf einem gemeinsamen Umfangskreis am Rumpf 12. Die auf einem Umfangskreis am Rumpf 12 angeordneten Düsen 22 werden durch die Leitflossen 16, 14, 20 und 18 in vier Gruppen mit jeweils fünf Düsen unterteilt. Jeweils eine solche Gruppe von Düsen 22 ist dem I.,II.,III. und IV. Quadranten (Fig. 2) zugeordnet. Die Einteilung der Düsen 22 in den Reihen 26 und 28 ist entsprechend.A plurality of nozzles 22 are located on the fuselage 12 between the adjacent guide fins 14 and 16, with five nozzles 22 each being arranged one behind the other in a row running transversely to the direction of flight A of the missile 10 and three such rows 24, 26 and 28. Nozzles 22 arranged in this way are located between all the adjacent fins of the missile 10. All of the nozzles 22 arranged in a row are located on a common circumferential circle on the fuselage 12. The nozzles 22 arranged on a circumferential circle on the fuselage 12 are guided by the fins 16, 14, 20 and 18 in four groups of each divided into five nozzles. Such a group of nozzles 22 is the I., II., III. and IV. Quadrants (Fig. 2) assigned. The division of the nozzles 22 in the rows 26 and 28 is corresponding.

Über die Düsen 22 wird Brennstoff in die am Rumpf 12 entlangstreichende Luft eingespritzt. Sämtliche am Rumpf 12 angeordnete Düsen 22 sind über (nicht dargestellte) Leitungen mit einem Tank für Brennstoff (ebenfalls nicht dargestellt) verbunden. Je nachdem, in welcher Art und Weise die Flugrichtung des Flugkörpers 10 zu steuern ist, können entweder alle Düsen 22 eines Quadranten oder aber auch die Düsen mehrerer Quadranten mit Brennstoff versorgt werden. In jeder mit dem Tank verbundenen Leitung, über die sämtliche Düsen eines Quadranten mit Brennstoff versorgt werden, befindet sich ein Ventil zum Verschließen bzw. Öffnen dieser Leitung. Die Versorgung der Düsen 22 mit Brennstoff erfolgt also nach Quadranten selektiert. Zwischen der mittleren Reihe 26 und der letzten Reihe 28 (in Flugrichtung A betrachtet) befindet sich eine Zündvorrichtung 30 - beispielsweise in Form einer Zündkerze - zum Zünden des aus den Düsen 22 des betreffenden Quadranten austretenden Brennstoffs.Fuel is injected into the air flowing along the fuselage 12 via the nozzles 22. All of the nozzles 22 arranged on the fuselage 12 are connected to a tank for fuel (likewise not shown) via lines (not shown). Depending on the manner in which the flight direction of the missile 10 is to be controlled, either all the nozzles 22 of one quadrant or else the nozzles of several quadrants can be supplied with fuel. In each line connected to the tank, through which all nozzles of a quadrant are supplied with fuel, there is a valve for closing or opening this line. The supply of the nozzles 22 with fuel is therefore selected according to quadrants. Between the middle row 26 and the last row 28 (viewed in the direction of flight A) there is an ignition device 30 - for example in the form of a spark plug - for igniting the fuel emerging from the nozzles 22 of the quadrant in question.

Die Entstehung eines Rezirkulationsgebietes zwischen zwei benachbarten Leitflossen soll nachfolgend beispielhaft anhand des ersten Quadranten (anhand des Bereiches zwischen den Leitflossen 14 und 16) verdeutlicht werden. Bei Überschallgeschwindigkeit gehen von der Vorderkante 14F der Leitflosse 14 und von der Vorderkante 16F der Leitflosse 16 jeweils zwei Stoßwellenfronten aus. Dabei verläuft eine der beiden von der Vorderkante 14F der Leitflosse 14 ausgehenden Stoßwellenfronten in Richtung auf die Leitflosse 16, während eine der beiden von der Vorderkante 16F der Leitflosse 16 ausgehende Stoßwellenfront in Richtung auf die Leitflosse 14 verläuft. Die Stoßwellen jeder Stoßwellenfront werden an den entsprechenden Leitflossen reflektiert. Dabei interferieren die Stoßwellen mit der am Rumpf l2 des Flugkörpers 10 entlangstreichenden Luft. Die am Rumpf 12 entlangstreichende Luft ist aufgrund der Reibung am Rumpf 12 abgebremst, wodurch sich bei Interferenz mit den Stoßwellen Rezirkulationsgebiete bilden. Die aufgrund der Reibung am stärksten verlangsamte Strömung der am Rumpf entlangstreichenden Luft ergibt sich jeweils in der Ecke (Eckenströmung) zwischen den Leitflossen 14 bzw. 16 und dem Rumpf 12 des Flugkörpers 10. Daher ergeben sich auch die am stärksten ausgebildeten Rezirkulationsgebiete jeweils in der Nähe einer Leitflosse.The creation of a recirculation area between two adjacent guide fins is to be illustrated below using the first quadrant (using the area between guide fins 14 and 16) as an example. At supersonic speeds, two shock wave fronts originate from the leading edge 14F of the guide fin 14 and from the leading edge 16F of the guide fin 16. One of the two extends from the leading edge 14F of the guide fin 14 outgoing shock wave fronts in the direction of the guide fin 16, while one of the two shock wave fronts starting from the front edge 16F of the guide fin 16 runs in the direction of the guide fin 14. The shock waves of each shock wave front are reflected on the corresponding guide fins. The shock waves interfere with the air passing along the fuselage l2 of the missile 10. The air flowing along the fuselage 12 is braked due to the friction on the fuselage 12, as a result of which recirculation areas are formed in the event of interference with the shock waves. The most slowed down flow of the air flowing along the fuselage occurs in the corner (corner flow) between the guide fins 14 and 16 and the fuselage 12 of the missile 10. Therefore, the most highly developed recirculation areas are also in the vicinity a guide fin.

Für den in Fig. 1 dargestellten Flugkörper 10 sei angenommen, daß die Reflektionen der Stoßwellen einer Stoßwellenfront an den Leitflossen in einem Bereich erfolgen, der sich zwischen der vordersten Reihe 24 und der hintersten Reihe 28 an den Leitflossen erstreckt. Die am stärksten ausgebildeten Rezirkulationsgebiete ergeben sich, wie bereits erwähnt, in unmittelbarer Nähe der Leitflossen. Kleinere Rezirkulationsgebiete entstehen in dem Bereich am Rumpf, in dem die mittleren Düsen der einzelnen Reihen angeordnet sind, da auch hier eine Interferenz zwischen den Stoßwellen der Stoßwellenfront und der am Rumpf 12 entlangstreichenden Luft erfolgt.For the missile 10 shown in FIG. 1, it is assumed that the reflections of the shock waves of a shock wave front occur on the guide fins in a region which extends between the foremost row 24 and the rearmost row 28 on the guide fins. The most highly developed recirculation areas are, as already mentioned, in the immediate vicinity of the guide fins. Smaller recirculation areas arise in the area on the fuselage in which the middle nozzles of the individual rows are arranged, since here too there is interference between the shock waves of the shock wave front and the air passing along the fuselage 12.

Zur Steuerung des Flugkörpers 10 wird je nach Bedarf über die Düsen 22 eines oder mehrerer Quadranten Brennstoff in die am Rumpf 12 entlangstreichende Luft eingespritzt. Der austretende Brennstoff wird mit Hilfe der Zündvorrichtung 30 gezündet, wobei sich eine lokal stabile Flamme bildet. Die Verbrennung des Brennstoffs verursacht eine Volumenvergrößerung des zwischen den Leitflossen 14 und 16 am Rumpf 12 befindlichen Gemisches aus verbranntem Brennstoff und am Rumpf 12 entlangstreichende Luft. Diese Volumenvergrößerung hat einen Druckanstieg in dem durch die Leitflossen 14 und 16 genau begrenzten Gebiet zur Folge. Der in diesem Gebiet erhöhte Druck wirkt auf den Rumpf 12, wodurch eine Querkraft, die quer zur Flugrichtung A des Flugkörpers 10 gerichtet ist, erzeugt wird. Die Stärke der Querkraft kann über die durch die Düsen 22 eines Quadranten austretende Brennstoffmenge pro Zeiteinheit reguliert werden.To control the missile 10, fuel is injected via the nozzles 22 of one or more quadrants as required injected into the air flowing along the fuselage 12. The emerging fuel is ignited with the aid of the ignition device 30, a locally stable flame being formed. The combustion of the fuel causes an increase in the volume of the mixture of burned fuel located on the fuselage 12 between the guide fins 14 and 16 and air flowing along the fuselage 12. This increase in volume results in an increase in pressure in the area precisely delimited by the guide fins 14 and 16. The increased pressure in this area acts on the fuselage 12, as a result of which a transverse force which is directed transversely to the flight direction A of the missile 10 is generated. The strength of the transverse force can be regulated via the amount of fuel exiting through the nozzles 22 of a quadrant per unit of time.

Der Druck des aus den Düsen 22 austretenden gasförmigen oder flüssigen Brennstoffs ist nur so groß, daß er ausreicht, um den Brennstoff aus dem Rumpf l2 austreten zu lassen. Durch diesen Druck allein wird dem Flugkörper kein wesentlicher Steuerungsimpuls erteilt. Wird als Brennstoff Wasserstoff verwendet, so erfolgt ab einer bestimmten Geschwindigkeit des Flugkörpers 10 eine Selbstentzündung des Wasserstoffs aufgrund der hohen Stautemperatur der am Rumpf 12 entlangstreichenden Luft. Die Zündungstemperatur für Wasserstoff liegt bei ca. 800°C. Wenn der Flugkörper 10 eine Geschwindigkeit größer als ca. 4 Mach bei Bodenbedingungen aufweist, ist die Temperatur der Luft am Flugkörper 12 aufgrund der hohen Stautemperaturen auf Werte größer als 800°C angestiegen, so daß eine sichere Selbstentzündung des Wasserstoffs erfolgt. Bei diesen Geschwindigkeitsbereichen des Flugkörpers können die in diesem ablaufenden Prozesse während der Steuerung dahingehend vereinfacht werden, daß bei jedem Steuerungsmanöver eine Ansteuerung der entsprechenden Zündvorrichtung nicht zu erfolgen braucht.The pressure of the gaseous or liquid fuel emerging from the nozzles 22 is only so great that it is sufficient to allow the fuel to emerge from the fuselage 12. This pressure alone does not give the missile any significant control impulse. If hydrogen is used as fuel, the hydrogen will self-ignite from a certain velocity of the missile 10 due to the high accumulation temperature of the air passing along the fuselage 12. The ignition temperature for hydrogen is around 800 ° C. If the missile 10 has a speed greater than approx. 4 Mach under ground conditions, the temperature of the air on the missile 12 has risen to values greater than 800 ° C. due to the high accumulation temperatures, so that the hydrogen ignites reliably. At these speed ranges of the missile, those running in it can Processes during the control are simplified in such a way that the control of the corresponding ignition device need not take place with each control maneuver.

Claims (7)

  1. Missile, in particular a supersonic missile, having a body (12) and, as seen in flight direction (A) at the rear body end, a number of tail fins (14,16,18,20) and, on the body (12), between the tail fins (14,16,18,20), a plurality of outflow apertures (22) for changing the flight direction (A),
    characterized in
    that ignitable fuel is ejected from the outflow apertures (22) for changing the flight direction (A) and that the outflow apertures (22) are arranged in areas where, due to shock waves originating from the front edges (14F,16F,18F,20F) of the tail fins (14,16,18,20), recirculation of the air passing along the body (12) is generated.
  2. Missile according to claim 1, characterized in that both sides of the front edges (14F,16F,18F,20F) of the tail fins are of a sharpened acute-angled form.
  3. Missile according to claim 1 or 2, characterized in that between two adjacent tail fins, an ignition means (30) is provided for the ignition of fuel ejected from the outflow aperture (22).
  4. Missile according to any one of claims 1 to 3, characterized in that between adjacent tail fins (14,16, 18,20), respectively, a plurality of outflow apertures (22) are arranged in a row (26), the outer outflow apertures of a row being disposed respectively in the direct vicinity of a tail fin (14,16,18,20).
  5. Missile according to claim 4, characterized in that the outflow apertures (22) are situated on a sole circumferential circle at the body (12).
  6. Missile according to claim 4 or 5, characterized in that a plurality of successive rows (24,26,28) of outflow apertures (22) are provided.
  7. Missile according to any one of the preceding claims, characterized in that the used fuel is hydrogen.
EP86117113A 1985-12-28 1986-12-09 Missile Expired - Lifetime EP0238724B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE3546269 1985-12-28
DE3546269A DE3546269C1 (en) 1985-12-28 1985-12-28 Missile

Publications (2)

Publication Number Publication Date
EP0238724A1 EP0238724A1 (en) 1987-09-30
EP0238724B1 true EP0238724B1 (en) 1991-05-15

Family

ID=6289654

Family Applications (1)

Application Number Title Priority Date Filing Date
EP86117113A Expired - Lifetime EP0238724B1 (en) 1985-12-28 1986-12-09 Missile

Country Status (4)

Country Link
US (1) US4712748A (en)
EP (1) EP0238724B1 (en)
DE (1) DE3546269C1 (en)
IL (1) IL81005A0 (en)

Families Citing this family (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3804931A1 (en) * 1988-02-17 1989-08-31 Deutsch Franz Forsch Inst Method for directional control of a missile flying in the relatively high supersonic domain, and such a missile
DE3937743A1 (en) * 1989-11-13 1991-05-16 Deutsch Franz Forsch Inst Supersonic missile with fuel ejector nozzle - has projecting rods facilitating flight control
US5070761A (en) * 1990-08-07 1991-12-10 The United States Of America As Represented By The Secretary Of The Navy Venting apparatus for controlling missile underwater trajectory
FR2684723B1 (en) * 1991-12-10 1995-05-19 Thomson Csf SOLID PROPERGOL PROPELLER WITH MODULAR PUSH AND MISSILE EQUIPPED.
US5318256A (en) * 1992-10-05 1994-06-07 Rockwell International Corporation Rocket deceleration system
US6178741B1 (en) * 1998-10-16 2001-01-30 Trw Inc. Mems synthesized divert propulsion system
US6752351B2 (en) * 2002-11-04 2004-06-22 The United States Of America As Represented By The Secretary Of The Navy Low mass flow reaction jet
US7416154B2 (en) * 2005-09-16 2008-08-26 The United States Of America As Represented By The Secretary Of The Army Trajectory correction kit
DE102005052474B3 (en) * 2005-11-03 2007-07-12 Junghans Feinwerktechnik Gmbh & Co. Kg Spiked artillery projectile
US8618455B2 (en) * 2009-06-05 2013-12-31 Safariland, Llc Adjustable range munition
CN106202807B (en) * 2016-07-22 2019-06-18 北京临近空间飞行器***工程研究所 Differentiate the method for space flight body portion shock wave/leading edge class Shock wave interaction occurrence condition and type

Family Cites Families (8)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3304029A (en) * 1963-12-20 1967-02-14 Chrysler Corp Missile directional control system
US3282541A (en) * 1965-02-19 1966-11-01 James E Webb Attitude control system for sounding rockets
US3749334A (en) * 1966-04-04 1973-07-31 Us Army Attitude compensating missile system
US3637167A (en) * 1969-11-05 1972-01-25 Mc Donnell Douglas Corp Missile steering system
DE2809281C2 (en) * 1978-03-03 1984-01-05 Emile Jean Versailles Stauff Control device for a self-rotating projectile
DE2846372C2 (en) * 1978-10-25 1985-11-21 Rheinmetall GmbH, 4000 Düsseldorf Projectile with radially directed control nozzles for final phase control
DE3378783D1 (en) * 1983-01-20 1989-02-02 Ford Aerospace & Communication Ram air combustion steering system for a guided missile
DE3340037A1 (en) * 1983-11-05 1985-05-23 Diehl GmbH & Co, 8500 Nürnberg ACTUATING SYSTEM FOR STEERED MISSIONS FLYING WITH SUPERVISOR SPEED

Also Published As

Publication number Publication date
IL81005A0 (en) 1987-03-31
EP0238724A1 (en) 1987-09-30
US4712748A (en) 1987-12-15
DE3546269C1 (en) 1987-08-13

Similar Documents

Publication Publication Date Title
EP0238724B1 (en) Missile
DE3004047C2 (en) Armor-piercing projectile
DE2214802A1 (en) SOLID ROCKET ENGINE
DE2452053A1 (en) DEVICE FOR LAUNCHING ROCKET-PROPELLED AIRCRAFT
DE3340037C2 (en)
DE2846372C2 (en) Projectile with radially directed control nozzles for final phase control
DE2027371C3 (en) Missile with stabilizing wings and a solid fuel engine with two separate propellants
DE69605579T2 (en) Projectile control with a multi-chamber, single jet engine
DE2160324C2 (en) Missile with deployable stabilization surfaces
DE3804931C2 (en)
DE1157929B (en) Aircraft with jet engines that are equipped with thrust reversers
DE1122776B (en) Flame holder unit
DE3804930C2 (en)
DE2757664A1 (en) DEVICE FOR CHANGING THE AIRWAY OF ONE FLOOR
DE2156974A1 (en) Container for scatter weapons for air deployment - is attached to carrying vehicle, steerable by remote control and has horizontal discharge tubes
DE3727532C2 (en) Tandem floor
DE2055707C1 (en) Launcher for a rocket-propelled missile
DE2856033C2 (en) Device for controlling and stabilizing an aircraft
DD301649A7 (en) PANZERABWEHRRAKETE WITH OUTBOARD AND MARINE ENGINE
DE1453854C1 (en) Remote-controlled missile
DE69001532T2 (en) Device for braking bombs after being dropped from an aircraft.
DE19509346C2 (en) Tail stabilized missile
DE2726551C2 (en)
DE2428402C3 (en) Launch device for rocket-propelled reverse runner missiles
DE2335126A1 (en) DEFLECTION DEVICE FOR ROCKET-PROPELLED BULLETS

Legal Events

Date Code Title Description
PUAI Public reference made under article 153(3) epc to a published international application that has entered the european phase

Free format text: ORIGINAL CODE: 0009012

AK Designated contracting states

Kind code of ref document: A1

Designated state(s): FR GB SE

17P Request for examination filed

Effective date: 19871017

17Q First examination report despatched

Effective date: 19890621

RAP1 Party data changed (applicant data changed or rights of an application transferred)

Owner name: DEUTSCHE FORSCHUNGSANSTALT FUER LUFT- UND RAUMFAHR

GRAA (expected) grant

Free format text: ORIGINAL CODE: 0009210

AK Designated contracting states

Kind code of ref document: B1

Designated state(s): FR GB SE

GBT Gb: translation of ep patent filed (gb section 77(6)(a)/1977)
ET Fr: translation filed
PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: GB

Effective date: 19911209

PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: SE

Effective date: 19911210

PLBE No opposition filed within time limit

Free format text: ORIGINAL CODE: 0009261

STAA Information on the status of an ep patent application or granted ep patent

Free format text: STATUS: NO OPPOSITION FILED WITHIN TIME LIMIT

26N No opposition filed
GBPC Gb: european patent ceased through non-payment of renewal fee
PG25 Lapsed in a contracting state [announced via postgrant information from national office to epo]

Ref country code: FR

Effective date: 19920831

REG Reference to a national code

Ref country code: FR

Ref legal event code: ST

EUG Se: european patent has lapsed

Ref document number: 86117113.0

Effective date: 19920704