EP0238724A1 - Missile - Google Patents

Missile Download PDF

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Publication number
EP0238724A1
EP0238724A1 EP86117113A EP86117113A EP0238724A1 EP 0238724 A1 EP0238724 A1 EP 0238724A1 EP 86117113 A EP86117113 A EP 86117113A EP 86117113 A EP86117113 A EP 86117113A EP 0238724 A1 EP0238724 A1 EP 0238724A1
Authority
EP
European Patent Office
Prior art keywords
missile
fuselage
guide fins
fuel
fins
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP86117113A
Other languages
German (de)
English (en)
Other versions
EP0238724B1 (fr
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/fr
Application granted granted Critical
Publication of EP0238724B1 publication Critical patent/EP0238724B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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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
    • F42B10/66—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 and a plurality of guide fins at its rear end in the direction of flight.
  • Missiles usually have fins on the fuselage.
  • Such missiles such as missiles or cruise missiles, are generally controlled by oars, the position of which is changed relative to the missile for the purpose of control.
  • Another control option is a so-called thrust vector control, in which the propellant gas stream emerging from the missile is deflected.
  • Another control option is to allow control gases to emerge essentially transversely and in pulses from the fuselage's fuselage.
  • a missile controlled by such a pulse control has openings at the rear end of the fuselage in the direction of flight through which a control gas emerges, with the Missile a force is applied transversely to the direction of flight. Depending on the desired direction of flight, the gas only emerges from certain holes.
  • the suddenly escaping gas gives the missile an impulse that changes the direction of flight.
  • the gas must be forced through the holes into the air flowing past the fuselage at high pressure in order to achieve an effective control effect.
  • relatively large, high-pressure gas masses are required to control the missile, which leads to weight and space problems in the missile.
  • the invention has for its object to provide a missile in which only small amounts of control gas are required to control its flight direction.
  • At least one outflow opening is arranged on the fuselage between adjacent guide fins in an area in which the speed of the air flowing along the fuselage is reduced as a result of shock waves emanating from the leading edges of the guide fins, from which for changing the Direction of flight fuel emerges.
  • At least one outflow opening is arranged on the fuselage between adjacent guide fins.
  • a fuel emerges from the outflow opening and ignites outside the missile. 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.
  • Such an 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.
  • 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
  • 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 increase in pressure, 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 such as rudders 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 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 formed 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.
  • nozzles 22 Between the adjacent guide fins 14 and 16 there are a plurality of nozzles 22 on the fuselage 12, five nozzles 22 each being arranged in a row running transversely to the flight direction A of the missile 10 and three such rows 24, 26 and 28 one behind the other. 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, each with divided into five nozzles. Such a group of nozzles 22 is the I., II., III. and IV. Quadrants (Fig. 2) assigned. The arrangement of the nozzles 22 in the rows 26 and 28 is corresponding.
  • Fuel is injected via the nozzles 22 into the air flowing along the fuselage 12. 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.
  • 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 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 strongly developed recirculation areas also result in the vicinity a guide fin.
  • one or more quadrants are used to fire the nozzles 22 as required fabric is 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 volume of the hull located between the guide fins 12 mixture 14 and 16 of burnt fuel and sweeping along the hull 12 air.
  • 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 exit 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, they can decrease in it current 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)
EP86117113A 1985-12-28 1986-12-09 Missile Expired - Lifetime EP0238724B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE3546269 1985-12-28
DE3546269A DE3546269C1 (de) 1985-12-28 1985-12-28 Flugkoerper

Publications (2)

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

Family

ID=6289654

Family Applications (1)

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

Country Status (4)

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

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3937743A1 (de) * 1989-11-13 1991-05-16 Deutsch Franz Forsch Inst Flugkoerper
FR2684723A1 (fr) * 1991-12-10 1993-06-11 Thomson Csf Propulseur a propergol solide a poussee modulable et missile equipe.

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3804931A1 (de) * 1988-02-17 1989-08-31 Deutsch Franz Forsch Inst Verfahren zur richtungssteuerung eines im hoeheren ueberschallbereich fliegenden flugkoerpers und derartiger flugkoerper
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
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 (de) * 2005-11-03 2007-07-12 Junghans Feinwerktechnik Gmbh & Co. Kg Drallstbilisiertes Artillerieprojektil
US8618455B2 (en) * 2009-06-05 2013-12-31 Safariland, Llc Adjustable range munition
CN106202807B (zh) * 2016-07-22 2019-06-18 北京临近空间飞行器系统工程研究所 判别航天器身部激波/前缘类激波干扰发生条件及类型的方法

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3749334A (en) * 1966-04-04 1973-07-31 Us Army Attitude compensating missile system
DE2846372A1 (de) * 1978-10-25 1983-10-20 Rheinmetall GmbH, 4000 Düsseldorf Verfahren und vorrichtung zur steigerung der treffgenauigkeit von geschossen
WO1984002975A1 (fr) * 1983-01-20 1984-08-02 Ford Aerospace & Communication Systeme de guidage par combustion d'air sous pression dynamique
DE3340037A1 (de) * 1983-11-05 1985-05-23 Diehl GmbH & Co, 8500 Nürnberg Stellsystem fuer gelenkte, mit ueberschallgeschwindigkeit fliegende flugkoerper

Family Cites Families (4)

* 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
US3637167A (en) * 1969-11-05 1972-01-25 Mc Donnell Douglas Corp Missile steering system
DE2809281C2 (de) * 1978-03-03 1984-01-05 Emile Jean Versailles Stauff Steuervorrichtung für ein Geschoß mit Eigendrehung

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3749334A (en) * 1966-04-04 1973-07-31 Us Army Attitude compensating missile system
DE2846372A1 (de) * 1978-10-25 1983-10-20 Rheinmetall GmbH, 4000 Düsseldorf Verfahren und vorrichtung zur steigerung der treffgenauigkeit von geschossen
WO1984002975A1 (fr) * 1983-01-20 1984-08-02 Ford Aerospace & Communication Systeme de guidage par combustion d'air sous pression dynamique
DE3340037A1 (de) * 1983-11-05 1985-05-23 Diehl GmbH & Co, 8500 Nürnberg Stellsystem fuer gelenkte, mit ueberschallgeschwindigkeit fliegende flugkoerper

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE3937743A1 (de) * 1989-11-13 1991-05-16 Deutsch Franz Forsch Inst Flugkoerper
FR2684723A1 (fr) * 1991-12-10 1993-06-11 Thomson Csf Propulseur a propergol solide a poussee modulable et missile equipe.

Also Published As

Publication number Publication date
DE3546269C1 (de) 1987-08-13
US4712748A (en) 1987-12-15
EP0238724B1 (fr) 1991-05-15
IL81005A0 (en) 1987-03-31

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