EP0433128A1 - Überschallgeschoss mit Lenkung mittels paarweise wirkenden Luftbremsen - Google Patents

Überschallgeschoss mit Lenkung mittels paarweise wirkenden Luftbremsen Download PDF

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Publication number
EP0433128A1
EP0433128A1 EP90403430A EP90403430A EP0433128A1 EP 0433128 A1 EP0433128 A1 EP 0433128A1 EP 90403430 A EP90403430 A EP 90403430A EP 90403430 A EP90403430 A EP 90403430A EP 0433128 A1 EP0433128 A1 EP 0433128A1
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EP
European Patent Office
Prior art keywords
fuselage
missile
flap
missile according
distance
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
EP90403430A
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English (en)
French (fr)
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EP0433128B1 (de
Inventor
Léon Boyadjian
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Airbus Group SAS
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Airbus Group SAS
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Publication of EP0433128A1 publication Critical patent/EP0433128A1/de
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Publication of EP0433128B1 publication Critical patent/EP0433128B1/de
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    • 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

Definitions

  • the invention relates to the guidance of supersonic missiles (or sub-missiles), in particular in the non-propelled phase. It targets in particular, but not exclusively, missiles powered and guided at high speed (at least Mach 2, in practice Mach 4 to 5) of the type called "hyperveloces", operating at low altitude and intended for the neutralization of air attackers or terrestrial discovered late, such as for example armed helicopters, fast planes evolving at low altitude and capable of sudden fallback (evasive) reactions or tanks.
  • the invention thus relates in particular to a missile whose mission comprises a first phase of rapid acceleration during which the abscissa relative to the missile of its center of gravity varies much longitudinally by consumption of propellant, then a second phase of deceleration during which this abscissa of the center of gravity remains fixed.
  • the invention also relates to an inert missile (or sub-missile or projectile) previously accelerated to the desired speed by means of propulsion dissociated after acceleration.
  • inert missile or sub-missile or projectile
  • propulsion dissociated after acceleration we find in this case the aforementioned phase with a fixed center of gravity with respect to the missile.
  • the object of the invention is to overcome the aforementioned drawbacks, in particular in the guiding phase without propulsion, by means of the combination of one or more retractable deflector flap (s), or spoilers, with a fixed wing (including possible "ducks"), which leads to a significant dynamic pressure effect corresponding to the exit of the spoiler.
  • missile In the context of the invention, the term "missile" will be interpreted in a broad sense, including the concepts of missiles proper, sub-missiles, and projectiles.
  • the invention proposes a guided supersonic missile, comprising a fuselage terminated at the front by a nose and at the rear by a base, and externally provided with fixed rear stabilizers, characterized in that it comprises, longitudinally at a distance from the center of gravity at least one movable flap, called a spoiler, between a configuration erased inside the fuselage and an active configuration deployed transversely in which this flap projects laterally with respect to this fuselage.
  • Such a missile lends itself to piloting by pitch and / or yaw torque which makes it possible, in response to a spoiler output command, to obtain a very rapidly high load factor for a hyperveloce missile evolving at low altitude.
  • This command is advantageously progressive (even proportional) so as to generate the effect necessary but just sufficient to pilot the hyperveloce missile.
  • the invention thus proposes the addition, to the wing and of any fixed ducks, of front or rear spoilers with proportional controls.
  • the addition of the ducks is very beneficial: the center of thrust of the interactions is located well in front of the spoiler which gives a much higher pitch up moment.
  • the effect of the rear spoiler is of the same order of magnitude as that of the front spoiler with ducks, but the load factor is lower due to the resulting loss of rear lift.
  • the rear spoiler has the advantage of reducing the additional aerodynamic drag corresponding to its active position by more than half.
  • Figures 1, 4 and 5 show a missile 1 comprising a cylindrical fuselage 2 terminated at the front by a nose 3 formed by a warhead and terminated at the rear by a base 4 provided with four fixed tail 5 of trapezoidal plane shape .
  • this missile 1 is provided with four fixed tail units 6 or "ducks" of substantially trapezoidal flat shape running partly along the warhead 3 and partly of the fuselage.
  • this missile is conventional (apart from the torque piloting device described below) and will not be described further. It will only be recalled that, in the case of a propelled missile of the hyperveloce type, a propellant of any known known type equips the rear of this missile.
  • a movable transverse flap 7, or spoiler adapted to disappear in the space occupied by the missile and the warhead or, on the contrary, to deploy.
  • the spoiler is unique here. Its rear face is longitudinally at the same level as the rear edge of the ducks 6. This spoiler is here permanently disposed in a transverse plane in which it retracts or deploys.
  • FIGS. 2 and 6 show a missile 1 ′ similar to missile 1 (with reference numbers assigned the index "prime") except that it is devoid of any front duck.
  • Figures 3 and 7 show a missile 1 ⁇ similar to missile 1 (with reference numbers assigned the index "second") except that the spoiler 7 ⁇ is disposed at the rear near the base 4 ⁇ between two 5 ⁇ rear tail.
  • FIG. 8 visualizes the forces resulting from the exit of the spoiler 7 or 7 ′: there appears an axial component A of braking and a transverse component F L which reduces to the center of gravity at a torque M tending to raise the nose 3 of the missile, M ⁇ representing the upstream infinite Mach number.
  • FIG. 9 visualizes, for a fictitious piloting case according to the third of the four piloting concepts mentioned above, that is to say of the powered missile, the forces resulting from the encroachment of an interceptor of imaginary jet 9 which would intercept from the bottom the jet of thrust from the nozzle 8 that the missile has in its base: there appears an axial braking component A ′ oriented towards the front, and a transverse component F L ′ directed towards the bottom whose resultant P ′ is in the opposite direction to that of the case of FIG. 8, but which is reduced to the center of gravity with a torque in the same direction as in FIG. 8, M jet being the Mach number leaving the jet.
  • the load factor actually obtained is the sum of the aerodynamic load factor of the missile taking into account its instantaneous incidence and the load factor induced by the spoiler
  • the load factor actually obtained is worth the aerodynamic load factor of the missile minus the load factor induced by the spoiler. This explains the preference, from this point of view, for the front spoilers.
  • the aerodynamic characteristics of missiles 1, 1 ′ and 1 ⁇ were determined by wind tunnel tests carried out between Mach 1.6 and Mach 4.34 using reduced-scale models in accordance with Figures 1 to 3 of 41.4 mm diameter (caliber) and 585.6 mm long (i.e. an elongation of the new model - length / diameter ratio - of 14.14), and a circular meridian warhead and elongation 2.5.
  • the cylindrical fuselage was equipped at the base with four empennages of 142.6 mm wingspan with an apex located 533.6 mm from the front point of the nose.
  • the height of the spoiler, extended position, was 6.2 mm and its width of 26 mm so that it could be inserted between the front or rear empennages (ducks).
  • the front spoiler was located 2.5 calibers from the front tip while that of the rear spoiler was at the base of the missile at 13.8 calibers from the front tip, these spoilers protruding about 0, 15 caliber (about 1% of the fuselage length).
  • FIG. 11 presents three curves C1, C2 and C3 corresponding respectively to the configurations of FIGS. 1, 2 and 3. They reflect the correlation between n the load factor and M the Mach number.
  • Cm 104 mN and transposed to the actual dimensions of the missile taking into account the requested travel (approximately 26 mm on the scale of the missile); the arrangement described is that of the front spoiler in FIGS. 1 or 2.
  • the motor 10 is installed along a transverse axis of the missile, the movement of the motor being communicated to the spoiler 7 by a ball screw 11.
  • the references 12 and 13 designate pinions coupling the shaft 10A of the motor and the screw 11.
  • the reference 14 designates a nut housing integral with the spoiler.
  • the references 15 and 16 designate guide tabs for the spoiler and the reference 17 designates a displacement sensor.
  • the motor 20 is arranged along the axis of the missile.
  • the transmission of the movement is done by a rack 21, integral with the spoiler and meshing with a pinion 22 integral with the shaft 20A of the engine.
  • the spoiler is guided by tabs 23 and 24.
  • Zone 25 represents the electrical supply.
  • control is of the proportional type with an implantation of a displacement sensor (only shown in FIG. 12).
  • FIG. 16 shows an electric motor, controlling a pneumatic cylinder 31 acting on a lever 32 of fixed axis 33. This lever acts on a link 34 coupled to the spoiler, moreover guided by cleats 35 and 36.
  • the supply of said control can be either hot gas or cold gas (use of an on-board bottle).
  • the efforts and the response times of said envisaged solutions are compatible with the performance requested.
  • the conventional solution therefore has a weight balance which is approximately twice the balance proposed in the two aforementioned solutions proposed by the invention.
  • the missile can have commands in pitch and yaw achieved by the use of four spoilers located at the before.
  • front and rear spoilers are also to be considered as well as the combination of a control by front spoilers and a control by rear jets or vice versa.
  • the invention is not limited to the case of cylindrical fuselages, but also applies to the case of fuselages of polygonal section inscribed in a circle (square, octagon ...), or even of generally elliptical section, in particular inscribed in use ellipse (rectangle, rhombus ).
  • the aforementioned notion of "diameter” then designates an average transverse dimension.

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)
  • Toys (AREA)
  • Control Of Position, Course, Altitude, Or Attitude Of Moving Bodies (AREA)
EP90403430A 1989-12-12 1990-12-03 Überschallgeschoss mit Lenkung mittels paarweise wirkenden Luftbremsen Expired - Lifetime EP0433128B1 (de)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
FR8916406 1989-12-12
FR8916406A FR2655722B1 (fr) 1989-12-12 1989-12-12 Missile supersonique a pilotage en couple par spouilers.

Publications (2)

Publication Number Publication Date
EP0433128A1 true EP0433128A1 (de) 1991-06-19
EP0433128B1 EP0433128B1 (de) 1996-07-10

Family

ID=9388412

Family Applications (1)

Application Number Title Priority Date Filing Date
EP90403430A Expired - Lifetime EP0433128B1 (de) 1989-12-12 1990-12-03 Überschallgeschoss mit Lenkung mittels paarweise wirkenden Luftbremsen

Country Status (6)

Country Link
US (1) US5143320A (de)
EP (1) EP0433128B1 (de)
CA (1) CA2031283C (de)
DE (1) DE69027750T2 (de)
ES (1) ES2088999T3 (de)
FR (1) FR2655722B1 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE4239589A1 (de) * 1992-11-25 1994-05-26 Deutsche Aerospace Vorrichtung zum Steuern von Flugkörpern mit einem aerodynamisch wirkenden Steuerkörper
FR2702556A1 (fr) * 1993-03-08 1994-09-16 Giat Ind Sa Tête militaire incendiaire.

Families Citing this family (25)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5677508A (en) * 1995-08-15 1997-10-14 Hughes Missile Systems Company Missile having non-cylindrical propulsion section
GB9614133D0 (en) * 1996-07-05 1997-03-12 Secr Defence Means for increasing the drag on a munition
RU2126949C1 (ru) * 1997-04-29 1999-02-27 Конструкторское бюро приборостроения Способ проверки приведения в рабочее состояние герметичного автопилотного блока
DE19839493C1 (de) * 1998-08-29 1999-12-30 Inst Franco Allemand De Rech D Verfahren zur Steuerung eines Überschallflugkörpers sowie Überschallflugkörper
RU2186332C2 (ru) * 2000-06-26 2002-07-27 Открытое акционерное общество АК "Туламашзавод" Управляемая ракета
RU2184342C2 (ru) * 2000-08-15 2002-06-27 Государственное унитарное предприятие "Конструкторское бюро приборостроения" Аэродинамический руль (варианты)
WO2003064254A2 (en) 2002-01-30 2003-08-07 Gulfstream Aerospace Corporation Fuselage shaping and inclusion of spike on a supersonic aircraft for controlling and reducing sonic boom
US6698684B1 (en) 2002-01-30 2004-03-02 Gulfstream Aerospace Corporation Supersonic aircraft with spike for controlling and reducing sonic boom
US6629668B1 (en) * 2002-02-04 2003-10-07 The United States Of America As Represented By The Secretary Of The Army Jump correcting projectile system
US6981672B2 (en) * 2003-09-17 2006-01-03 Aleiant Techsystems Inc. Fixed canard 2-D guidance of artillery projectiles
US7367530B2 (en) * 2005-06-21 2008-05-06 The Boeing Company Aerospace vehicle yaw generating systems and associated methods
US8333076B2 (en) * 2005-12-15 2012-12-18 Gulfstream Aerospace Corporation Isentropic compression inlet for supersonic aircraft
US7611095B1 (en) * 2006-04-28 2009-11-03 The Boeing Company Aerodynamic re-entry vehicle control with active and passive yaw flaps
US8393158B2 (en) 2007-10-24 2013-03-12 Gulfstream Aerospace Corporation Low shock strength inlet
ES2656306T3 (es) * 2008-05-30 2018-02-26 Saab Ab Disposición y procedimiento de lanzamiento de contramedidas
US8502126B2 (en) * 2010-05-27 2013-08-06 Raytheon Company System and method for navigating an object
RU2438095C1 (ru) * 2010-06-21 2011-12-27 Государственное унитарное предприятие "Конструкторское бюро приборостроения" Управляемый вращающийся снаряд
US8525090B1 (en) * 2010-06-23 2013-09-03 The United States Of America As Represented By The Secretary Of The Army Pneumatically actuated control surface for airframe body
RU2580376C2 (ru) * 2014-07-29 2016-04-10 Николай Евгеньевич Староверов Крылатая ракета, в частности - противокорабельная (варианты)
RU2690236C1 (ru) * 2018-04-03 2019-05-31 Сергей Евгеньевич Угловский Сверхзвуковая вращающаяся ракета
DE102018005480A1 (de) * 2018-07-11 2020-01-16 Mbda Deutschland Gmbh Flugkörper
RU2703017C1 (ru) * 2018-09-24 2019-10-15 Сергей Евгеньевич Угловский Сверхзвуковая вращающаяся ракета
CN109823515B (zh) * 2019-01-24 2020-12-15 北京理工大学 设置在制导飞行器上的扰流板系统及应用其的方法
CN113830290B (zh) * 2021-09-03 2022-04-26 中国空气动力研究与发展中心低速空气动力研究所 一种伸缩式涡流发生器及其构成的桨毂
CN114486159A (zh) * 2021-12-30 2022-05-13 中国航天空气动力技术研究院 内埋武器机弹分离相容性前缘锯齿扰流板控制及验证方法

Citations (6)

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Publication number Priority date Publication date Assignee Title
US3305194A (en) * 1960-03-08 1967-02-21 Robert G Conard Wind-insensitive missile
GB1188651A (en) * 1962-03-05 1970-04-22 British Aircraft Corp Ltd Improvements in or relating to Missiles
FR2342201A1 (fr) * 1976-02-26 1977-09-23 Hawker Siddeley Dynamics Ltd Regulation aerodynamique d'un vehicule au moyen de gouvernes auxiliaires escamotables
EP0013096A1 (de) * 1978-12-29 1980-07-09 The Commonwealth Of Australia Anordnung für ausklappbare Flügelflächen
US4497460A (en) * 1983-03-25 1985-02-05 The United States Of America As Represented By The Secretary Of The Navy Erodale spin turbine for tube-launched missiles
EP0264529A1 (de) * 1986-08-19 1988-04-27 Rheinmetall GmbH Flugkörper mit Stabilisierungseinrichtung

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FR496912A (fr) * 1916-08-08 1919-11-20 Charles Leopold Mayer Procédé pour faire du tir plongeant sans diminuer la charge
US2922600A (en) * 1956-04-18 1960-01-26 John B Craft Automatic guidance system
US3136250A (en) * 1962-05-04 1964-06-09 Samuel A Humphrey Integrated auxiliary power unit
US3188958A (en) * 1963-03-11 1965-06-15 James D Burke Range control for a ballistic missile
CH480613A (de) * 1967-09-11 1969-10-31 Oerlikon Buehrle Ag Geschoss mit Bremsflügeln
US3759466A (en) * 1972-01-10 1973-09-18 Us Army Cruise control for non-ballistic missiles by a special arrangement of spoilers
US4327884A (en) * 1980-01-23 1982-05-04 The United States Of America As Represented By The Secretary Of The Air Force Advanced air-to-surface weapon
US4699333A (en) * 1984-11-07 1987-10-13 The Boeing Company On-board flight control panel system

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3305194A (en) * 1960-03-08 1967-02-21 Robert G Conard Wind-insensitive missile
GB1188651A (en) * 1962-03-05 1970-04-22 British Aircraft Corp Ltd Improvements in or relating to Missiles
FR2342201A1 (fr) * 1976-02-26 1977-09-23 Hawker Siddeley Dynamics Ltd Regulation aerodynamique d'un vehicule au moyen de gouvernes auxiliaires escamotables
EP0013096A1 (de) * 1978-12-29 1980-07-09 The Commonwealth Of Australia Anordnung für ausklappbare Flügelflächen
US4497460A (en) * 1983-03-25 1985-02-05 The United States Of America As Represented By The Secretary Of The Navy Erodale spin turbine for tube-launched missiles
EP0264529A1 (de) * 1986-08-19 1988-04-27 Rheinmetall GmbH Flugkörper mit Stabilisierungseinrichtung

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE4239589A1 (de) * 1992-11-25 1994-05-26 Deutsche Aerospace Vorrichtung zum Steuern von Flugkörpern mit einem aerodynamisch wirkenden Steuerkörper
FR2702556A1 (fr) * 1993-03-08 1994-09-16 Giat Ind Sa Tête militaire incendiaire.

Also Published As

Publication number Publication date
US5143320A (en) 1992-09-01
DE69027750T2 (de) 1996-11-28
DE69027750D1 (de) 1996-08-14
EP0433128B1 (de) 1996-07-10
FR2655722B1 (fr) 1992-03-13
CA2031283A1 (en) 1991-06-13
ES2088999T3 (es) 1996-10-01
CA2031283C (en) 2001-05-29
FR2655722A1 (fr) 1991-06-14

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