EP0918976B1 - Procédé pour la fabrication des missiles ou des composants de missile - Google Patents

Procédé pour la fabrication des missiles ou des composants de missile Download PDF

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Publication number
EP0918976B1
EP0918976B1 EP97935567A EP97935567A EP0918976B1 EP 0918976 B1 EP0918976 B1 EP 0918976B1 EP 97935567 A EP97935567 A EP 97935567A EP 97935567 A EP97935567 A EP 97935567A EP 0918976 B1 EP0918976 B1 EP 0918976B1
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EP
European Patent Office
Prior art keywords
sic
missile
carbon
silicon carbide
reinforced
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
EP97935567A
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German (de)
English (en)
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EP0918976A1 (fr
Inventor
Manfred Braitinger
Manfred Selzer
Ulrich Papenburg
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.)
IABG Industrieanlagen Betriebs GmbH
Original Assignee
IABG Industrieanlagen Betriebs GmbH
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Application filed by IABG Industrieanlagen Betriebs GmbH filed Critical IABG Industrieanlagen Betriebs GmbH
Publication of EP0918976A1 publication Critical patent/EP0918976A1/fr
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Publication of EP0918976B1 publication Critical patent/EP0918976B1/fr
Anticipated expiration legal-status Critical
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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/34—Protection against overheating or radiation, e.g. heat shields; Additional cooling arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B12/00—Projectiles, missiles or mines characterised by the warhead, the intended effect, or the material
    • F42B12/72—Projectiles, missiles or mines characterised by the warhead, the intended effect, or the material characterised by the material

Definitions

  • the invention relates to a method for manufacturing of missiles or missile components.
  • the nose tip consists fixed fins or movable rudders or fins, thrusters, thrusters and nozzle neck inserts, Combustion chamber linings, rear cone, grille wing, fluidic elements and the radome from various metals and metal alloys.
  • these missile components are the most thermally and mechanically loaded Missile components.
  • the invention has for its object missiles or missile components such as ceramic nose tips, fixed fins or movable rudders, thrusters, thrusters and nozzle neck inserts, Combustion chamber linings, rear cone, grille wings, fluidic elements and radomes or subcomponents of these for missiles with high temperature, pressure and Abrasion resistance, erosion resistance, low density or low weight, high thermal conductivity, low thermal expansion with an almost unlimited To create a variety of geometries and shapes.
  • object missiles or missile components such as ceramic nose tips, fixed fins or movable rudders, thrusters, thrusters and nozzle neck inserts, Combustion chamber linings, rear cone, grille wings, fluidic elements and radomes or subcomponents of these for missiles with high temperature, pressure and Abrasion resistance, erosion resistance, low density or low weight, high thermal conductivity, low thermal expansion with an almost unlimited To create a variety of geometries and shapes.
  • Bow tip 1 fixed fins 2 or movable rudders [fins] 3, thrusters 4, thrusters or nozzle neck inserts 5, combustion chamber linings 6, rear cone 7, Lattice wings 8, fluidic elements 9 and radomes 10 or subcomponents of these consist of a fiber-reinforced ceramic or a combination of different ones fiber-reinforced ceramics and form a monolithic after infiltration Structure. Overall, the temperature resistance increases at the same time Weight reduction of these missile components.
  • C / SiC and / or C / C and / or SiC / SiC are excellent Strength properties up to high temperatures, which also use enable under severe conditions.
  • C / SiC and C / C and SiC / SiC with continuous fiber reinforcement as well as short fiber reinforced C / SiC and C / C and SiC / SiC.
  • the former material of C / SiC or C / C or SiC / SiC, which are laminated, pressed or wound can, is characterized by particularly high strength and particularly low density out.
  • Surface sealing can be used to increase the resistance to oxidation be worked.
  • Protective layers made of silicon carbide are preferred for this and / or silicon dioxide and / or molybdenum disilicide on the component surfaces upset. The latter is superfluous with short fiber reinforced C / SiC because of the material is particularly resistant to oxidation and corrosion.
  • Bow tips can be made from C / SiC blanks and / or C / C blanks 1, fixed fins 2 or movable rudders [fins] 3, thrusters 4, thrusters or nozzle neck inserts 5, combustion chamber linings 6, rear cone 7, grille wing 8, fluidic elements 9 and radomes 10 or partial components of these in any Geometry from one piece or from different individual segments by mechanical Machining can be easily shaped.
  • This construction is particularly suitable for C / SiC or C / C or SiC / SiC with short fiber reinforcement, the individual segments be mechanically processed before being combined or infiltrated.
  • Such a missile component 1-10 can also be easily made with fasteners such as.
  • Screws or bolts or flanges preferably made of C / SiC and / or C / C and / or SiC / SiC.
  • Cooling channels and / or recesses with round, rectangular or slit-shaped Cross section are introduced.
  • the method according to the invention provides for a design of the missile components 1-10 in hybrid and segment design.
  • raw bodies and sub-segments which are preferably made of C / SiC and / or C / C and / or SiC / SiC or from suitable combinations with continuous fiber reinforcement and / or short fiber reinforcement and by the subsequent Infiltration with silicon and / or silicon carbide and / or carbon of these individual segments monolithic missile components are designed in hybrid construction.
  • the inside wall of the missile or the thermal highly stressed areas of the missile can be suitably with C / SiC or C / C or SiC / SiC segments are lined and by cooling via cooling channels and / or with an insulation material, preferably made of C / SiC or C / C or SiC / SiC or from carbon fiber felts or graphite foil or combinations this is the temperature and pressure load of the metallic missile structure reduced as much as possible, provided and a monolithic missile component 1-10 can be put together.
  • the insulation materials can also under Interposition of spacers, preferably made of C / SiC or C / C or SiC / SiC, with the missile components 1-10 made of C / SiC and / or C / C with each other can be connected to give the desired monolithic structure.
  • the density and porosity of the C / SiC and / or the C / C and / or SiC / SiC material during infiltration or siliconization by the addition amount be adjusted to silicon, carbon or silicon carbide, so that C / SiC and / or C / C and / or SiC / SiC with high density and low porosity than thermomechanical support structure and / or lining and the C / SiC and / or C / C and / or SiC / SiC with low density or high porosity used as thermal insulation can be.
  • Density and porosity gradients over the Wall thickness of the missile components 1-10 can be set.
  • the missile component 1-10 is depending on the system used made from C / SiC and / or C / C and / or SiC / SiC individual segments, which then to a monolithic structure with carbon and / or silicon and / or silicon carbide are infiltrated and / or silicided together or the missile components 1-10 are manufactured in one piece, preferably by mechanical processing of a C / SiC and / or C / C and / or SiC / SiC blank.
  • C / SiC and / or C / C parts and / or SiC / SiC parts can also be the cooling channels (if necessary) or provide recesses to remove the heat.
  • the C / SiC and / or C / C body and / or SiC / SiC body 1-10 and the metallic Missile structures are equipped with suitable connecting elements such as Bolt-, Screw or flange connections, preferably made of C / SiC and / or C / C and / or SiC / SiC to connect with each other. Possibilities for this are in the pictures 2 to 9 shown.
  • Figure 1 shows a missile based on the current state of the art. Due to the high temperature and pressure loads, only heat-resistant metals and metal alloys with a high density are currently suitable, which have to be cooled due to their relatively low temperature resistance. In addition to these thermomechanical requirements, the metallic materials must also meet all requirements with regard to corrosion, machining, surface quality and weldability.
  • Figure 2 shows a nose tip 1 and a radome 10 of a missile.
  • the bow tip is particularly stressed by high pressures and high temperatures.
  • the weight of the nose tip can be reduced by at least 1 kg compared to a metallic nose tip.
  • Radomes are exposed to high pressures and high temperatures.
  • increased radar permeability and surface accuracy e.g. through grindability
  • construction of different wall thicknesses are required.
  • Figure 3 shows the stabilizing fins or fixed fins 2 and the tail cone 7 of a missile. Stresses on the fixed fins are mainly caused by high longitudinal and lateral acceleration forces and high temperatures.
  • the tail cone 7 of a missile is subjected to high pressures and high temperatures and serves to stabilize the missile.
  • the use of fiber-reinforced ceramics saves 3 kg in weight at the rear cone.
  • Figure 4 shows movable rudder or fin 3 and grid fins 8 are shown.
  • the movable rudders or fins 3 are subject to stresses caused by high longitudinal and lateral acceleration forces and high temperatures. They serve as an aerodynamic steering aid. Stresses due to high longitudinal and lateral acceleration forces and high temperatures also occur on the lattice wings 8. They serve both as an aerodynamic steering aid and to maintain the stability of the missile.
  • the grille wing looks like a narrow doormat attached to the tail of the missile, the openings of which are in the direction of flight and can be rotated about the longitudinal axis.
  • Fig. 5 the thrusters 4 according to the invention are shown.
  • a stress caused by high lateral forces, temperatures and abrasion by exhaust gases and solid particles (eg Al 2 O 3 particles) must be taken into account when designing thrusters.
  • the use of thrusters in the exhaust jet serves as an additional steering aid during the propulsion phase of the missile.
  • Jet rudders which are installed in the rear of a rocket nozzle in the exhaust gas jet for beam deflection, are subject to extremely high thermo-mechanical stresses due to the hot, reactive combustion gases and the high lateral forces.
  • Thermal shock resistance and good abrasion behavior against solid particles are additionally required for thrusters, since thrusters can suddenly be exposed to gas / particle flows at temperatures of 2500 ° C depending on the type of engine and type of fuel.
  • Figure 6 shows a thrust nozzle 5 and the typical embodiment of the combustion chamber lining 6 according to the invention.
  • the thruster is subjected to extremely high pressures and temperatures. Often have the engines of missiles for the individual thrust phases (ejection, Acceleration and marching phase) several and different number of thrusters.
  • Figure 7 shows typical fluidic elements 9 that are used as transverse thrust controls.
  • the method according to the invention provides that the thrust nozzle and / or the nozzle neck and / or the combustion chamber are lined with C / SiC and / or C / C segments and / or SiC / SiC segments.
  • the inner walls of the missiles are made of C / SiC and / or C / C and / or SiC / SiC individual segments.
  • the C / SiC and / or C / C and / or SiC / SiC segments are to be designed in such a way that the dividing slots, the gases under high pressure and high temperature, are not let through to the metallic missile structure.
  • the C / SiC and / or C / C parts and / or SiC / SiC parts can be adapted to the inner contour of the missile engine and thus enable a geometrical simplification of the missile structure.
  • the process provides that the C / SiC and / or C / C and / or SiC / SiC individual segments for the missile components (1-10) made of C / SiC and / or C / C and / or SiC / SiC blanks machined and before assembly in the Missile structure combined into a monolithic structure.
  • the cooling can optionally be introduced of cooling channels or recesses in the C / SiC and / or C / C and / or SiC / SiC structure or insulation with carbon felt or graphite foil or C / SiC or C / C or SiC / SiC or combinations of these.
  • the cooling with Depending on the requirements, the cooling channels can be configured in the missile structure on Transition metal to C / SiC and / or C / C and / or SiC / SiC or in C / SiC and / or C / C and / or SiC / SiC part itself. It is also a combination of the two Parts provided.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Physics & Mathematics (AREA)
  • Thermal Sciences (AREA)
  • Ceramic Products (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
  • Details Of Aerials (AREA)

Claims (3)

  1. Procédé pour la fabrication de missiles ou de composants de missiles soumis à de fortes sollicitations thermiques et mécaniques, comprenant les phases suivantes :
    a) fabrication d'ébauches en céramique crue renforcée par des fibres, à savoir carbure de silicium renforcé par des fibres de carbone (C/SiC) et/ou carbone renforcé par des fibres de carbone (C/C) et/ou carbure de silicium renforcé par des fibres de carbure de silicium (SiC/SiC) ;
    b) mise en forme et traitement mécanique des ébauches en céramique à l'état cru en fonction des géométries partielles du missile ou des composants du missile ;
    c) infiltration commune ou enrichissement commun au silicium des ébauches par du silicium et/ou carbure de silicium et/ou carbone pour obtenir une structure composite monolithique, les ébauches réunies formant la forme d'ensemble du missile ou des composants respectifs du missile.
  2. Procédé suivant la revendication 1, caractérisé en ce que la masse volumique et la porosité du matériau C/SiC et/ou C/C et/ou SiC/SiC sont réglées pendant la phase d'infiltration ou d'enrichissement au silicium par la quantité additionnelle de silicium, de carbone et/ou de carbure de silicium, de sorte qu'il en résulte une structure porteuse ayant d'une part une masse volumique élevée et une faible porosité comme partie porteuse et/ou revêtement, et ayant d'autre part une faible masse volumique et une porosité élevée comme isolation thermique.
  3. Procédé suivant la revendication 1, caractérisé en ce que des conduits de refroidissement sont pratiqués ou conformés par traitement mécanique dans les ébauches en céramique.
EP97935567A 1996-08-16 1997-08-04 Procédé pour la fabrication des missiles ou des composants de missile Expired - Lifetime EP0918976B1 (fr)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
DE19632893 1996-08-16
DE19632893A DE19632893C2 (de) 1996-08-16 1996-08-16 Verfahren zur Herstellung von Flugkörperkomponenten aus faserverstärkter Keramik
PCT/EP1997/004235 WO1998008044A1 (fr) 1996-08-16 1997-08-04 Composants de missile en ceramique renforcee par des fibres

Publications (2)

Publication Number Publication Date
EP0918976A1 EP0918976A1 (fr) 1999-06-02
EP0918976B1 true EP0918976B1 (fr) 2000-06-14

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ID=7802697

Family Applications (1)

Application Number Title Priority Date Filing Date
EP97935567A Expired - Lifetime EP0918976B1 (fr) 1996-08-16 1997-08-04 Procédé pour la fabrication des missiles ou des composants de missile

Country Status (5)

Country Link
US (1) US6460807B1 (fr)
EP (1) EP0918976B1 (fr)
AT (1) ATE193942T1 (fr)
DE (2) DE19632893C2 (fr)
WO (1) WO1998008044A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US7037602B2 (en) 2002-07-04 2006-05-02 Sgl Carbon Ag Multilayer composite

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US7800032B1 (en) * 2006-11-30 2010-09-21 Raytheon Company Detachable aerodynamic missile stabilizing system
US7829829B2 (en) * 2007-06-27 2010-11-09 Kazak Composites, Incorporated Grid fin control system for a fluid-borne object
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CN108007280B (zh) * 2017-12-28 2023-08-15 北京威标至远科技发展有限公司 一种舵机防热结构
GB2578572B (en) 2018-10-30 2022-08-17 Bae Systems Plc A sabot
CN114128045A (zh) * 2019-09-20 2022-03-01 阿塞尔桑电子工业及贸易股份公司 由不同组成的连续细丝制造多层陶瓷结构
EP4032144A4 (fr) * 2019-09-20 2022-11-16 Aselsan Elektronik Sanayi ve Ticaret Anonim Sirketi Procédé de fabrication de structures céramiques multicouches par des filaments continus de composition identique
CN112719804B (zh) * 2020-12-18 2022-06-07 湖北三江航天江北机械工程有限公司 一种训练用空空导弹吊挂组合的加工方法
CN112693623B (zh) * 2020-12-21 2022-05-27 中国空气动力研究与发展中心高速空气动力研究所 导弹栅格舵铰链力矩模型爪盘式自锁定位结构
CN112853250B (zh) * 2020-12-28 2022-08-05 哈尔滨工业大学 一种组合燃气舵构件的制备方法
CN114235321B (zh) * 2022-02-25 2022-04-26 中国空气动力研究与发展中心高速空气动力研究所 一种燃气舵和喷管一体化风洞测力实验装置

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* Cited by examiner, † Cited by third party
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US7037602B2 (en) 2002-07-04 2006-05-02 Sgl Carbon Ag Multilayer composite
DE10230231B4 (de) * 2002-07-04 2007-07-05 Sgl Carbon Ag Mehrschichtiger Verbundwerkstoff

Also Published As

Publication number Publication date
DE19632893C2 (de) 2001-02-08
WO1998008044A1 (fr) 1998-02-26
DE19632893A1 (de) 1998-02-19
DE59701892D1 (de) 2000-07-20
EP0918976A1 (fr) 1999-06-02
ATE193942T1 (de) 2000-06-15
US6460807B1 (en) 2002-10-08

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