EP0918976B1 - Process for manufacturing missiles or missile components - Google Patents

Process for manufacturing missiles or missile components 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
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EP97935567A
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German (de)
French (fr)
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EP0918976A1 (en
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Manfred Braitinger
Manfred Selzer
Ulrich Papenburg
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IABG Industrieanlagen Betriebs GmbH
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IABG Industrieanlagen Betriebs GmbH
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    • 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)
  • Physics & Mathematics (AREA)
  • Thermal Sciences (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Ceramic Products (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
  • Details Of Aerials (AREA)

Abstract

The missile comprises a nose, fixed fins or movable fins, gas rudders, propelling nozzles and blast pipe inserts, combustion chamber liners, tail cone, grid fins, fluid elements and radome or subcomponents of these made of carbon fiber-reinforced silicon carbide (C/SiC) and/or carbon fiber-reinforced carbon (C/C) and/or silicon carbide fiber-reinforced silicon carbide (SiC/SiC) which are integrated in a missile structure. The missile components may be made of C/SiC and/or C/C and/or SiC/SiC with continuous fiber reinforcement and/or chopped fiber reinforcement or combinations thereof. The missile components may be prepared by mechanical machining of C/SiC and/or C/C and/or SiC/SiC blanks either in a single piece or by co-infiltration with carbon or silicon or silicon carbide and/or co-siliconizing of separate C/SiC and/or C/C and/or SiC/SiC segments to result in a monolithic structure. Cooling of the missile components may be achieved by providing cooling ducts and/or recesses in the C/SiC and/or C/C structure and/or the SiC/SiC structure and/or by means of insulating materials such as carbon fiber felts or C/SiC or C/C or SiC/SiC or graphite sheets or combinations thereof.

Description

Die Erfindung betrifft ein Verfahren zur Herstellung von Flugkörpern oder Flugkörperkomponenten.The invention relates to a method for manufacturing of missiles or missile components.

Ein solches Verfahren ist aus der EP 0 541 917 A bekannt.Such a method is known from EP 0 541 917 A.

An Flugkörpern, die sich mit sehr hoher Geschwindigkeit in der bodennahen Atmosphäre bewegen, treten an exponierten Stellen, wie Kanten, Ecken und Spitzen wegen der aerodynamischen Aufheizung Oberflächentemperaturen von über 1700 °C auf. Sehr hohe Temperaturen von über 2500 °C treten an Bauteilen von Flugkörpermotoren auf, deren Festtreibstoffe teilweise mit Temperaturen von über 3500 °C verbrennen. Die betroffenen Bauteile sollen auch bei diesen Temperaturen noch über hinreichende Strukturfestigkeit und Funktionalität verfügen, um die Gesamtmission des Flugobjektes erfolgreich zu beenden. Bisher wurde die Strukturfestigkeit der meist metallischen Bauteile im Hochtemperatureinsatz durch Verwendung hochtemperaturbeständiger Metalle und Metallegierungen, Kühlung und thermische Isolierung realisiert. Diese Maßnahmen sind aufwendig, teuer und erfordern in allen Fällen zusätzliches Gewicht zur Erfüllung der Aufgabenstellung. Zusätzliches Gewicht ist bei mobilem Gerät, insbesondere bei Flugkörpern, nachteilig, so daß nach gewichtsreduzierten Lösungen gesucht werden muß. On missiles that move at very high speeds in the atmosphere close to the ground move, occur in exposed areas such as edges, corners and points the aerodynamic heating surface temperatures of over 1700 ° C on. Very high temperatures of over 2500 ° C occur on components of missile engines whose solid fuels sometimes reach temperatures of over 3500 ° C burn. The affected components should still be at these temperatures have sufficient structural strength and functionality to complete the mission to successfully complete the flying object. So far, the structural strength of mostly metallic components in high temperature use through use high temperature resistant metals and metal alloys, cooling and thermal Insulation realized. These measures are complex, expensive and require all Additional weight to fulfill the task. Additional weight is disadvantageous in the case of mobile devices, in particular missiles, so that after weight-reduced solutions must be sought.

Bei einer bekannten Ausführungsform eines Flugkörpers bestehen die Bugspitze, feste Flossen oder bewegliche Ruder bzw. Fins, Strahlruder, Schubdüsen und Düsenhalseinsätze, Brennkammerauskleidungen, Heckkonus, Gitterflügel, Fluidikelemente und das Radom aus verschiedenen Metallen und Metallegierungen. Dabei sind diese Flugkörperkomponenten die thermisch und mechanisch höchstbelasteten Bauteile des Flugkörpers.In a known embodiment of a missile, 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. Here these missile components are the most thermally and mechanically loaded Missile components.

Aufgrund der genannten hohen Temperaturen, hohen mechanischen Belastungen und hohen Drücke muß man bei der heutigen Auslegung dieser Flugkörperkomponenten hoch hitzebeständige Metalle oder Metallegierungen (z.B. Wolfram, Molybdän, Inconnel) mit hoher mechanischer Festigkeit und Temperaturbeständigkeit verwenden. Da diese temperaturbeständigen Metalle und Legierungen schon ab etwa 800 °C unter Festigkeitsverlust erweichen, muß zusätzlich aktiv gekühlt werden. Ein weiterer gravierender Nachteil der Flugkörperkomponenten aus Metall ist ihr hohes Gewicht, welches die Beschleunigung und Geschwindigkeit von Flugkörpern einschränkt.Because of the high temperatures mentioned, high mechanical loads and high pressures are required in today's design of these missile components highly heat-resistant metals or metal alloys (e.g. tungsten, molybdenum, Inconnel) with high mechanical strength and temperature resistance use. Since these temperature-resistant metals and alloys start from around Soften 800 ° C with loss of strength, must also be actively cooled. Another serious disadvantage of metal missile components is their high level Weight, which is the acceleration and velocity of missiles restricted.

Der Erfindung liegt die Aufgabe zugrunde, Flugkörper oder Flugkörperkomponenten wie keramische Bugspitzen, feste Flossen oder bewegliche Ruder [Fins], Strahlruder, Schubdüsen und Düsenhalseinsätze, Brennkammerauskleidungen, Heckkonus, Gitterflügel, Fluidikelemente und Radome oder Teilkomponenten aus diesen für Flugkörper mit hoher Temperatur-, Druck- und Abriebfestigkeit, Erosionsbeständigkeit, niedriger Dichte bzw. niedrigem Gewicht, hoher Wärmeleitfähigkeit, niedriger Wärmeausdehnung bei einer nahezu unbegrenzten Geometrie- und Formenvielfalt zu schaffen. 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.

Diese Aufgabe wird gelöst durch die Merkmale des Verfahrens Anspruchs 1.This task is solved by the characteristics of the Procedure claim 1.

Bugspitze 1, feste Flossen 2 oder bewegliche Ruder [Fins] 3, Strahlruder 4, Schubdüsen oder Düsenhalseinsätze 5, Brennkammerauskleidungen 6, Heckkonus 7, Gitterflügel 8, Fluidikelemente 9 und Radome 10 oder Teilkomponenten aus diesen bestehen also aus einer faserverstärkten Keramik oder aus Kombinationen verschiedener faserverstärkter Keramiken und bilden nach der Infiltration eine monolithische Struktur. Insgesamt erhöht sich die Temperaturbeständigkeit bei gleichzeitiger Gewichtsreduzierung dieser Flugkörperkomponenten.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.

Es wurde gefunden, daß C/SiC und/oder C/C und/oder SiC/SiC über hervorragende Festigkeitseigenschaften bis zu hohen Temperaturen verfügt, die einen Einsatz auch unter schweren Bedingungen ermöglichen. Hinzu kommt neben einer geringen Dichte hohe Verschleißfestigkeit, Oxidationsbeständigkeit sowie, neben der ausgezeichneten Temperaturbeständigkeit, eine hohe Temperaturwechselbeständigkeit. It has been found that 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. In addition to a minor Dense high wear resistance, resistance to oxidation and, in addition to the excellent Temperature resistance, high resistance to temperature changes.

Dabei ist es insbesondere bei Oberflächenversiegelung besonders gas- und flüssigkeitsdicht.It is particularly gastight and liquid-tight, especially with surface sealing.

Besonders hervorzuheben sind die große Geometrie- und Formenvielfalt bei gleichzeitig niedrigem Gewicht, sowie die hervorragende Temperaturfestigkeit und hohe bzw. einstellbare Wärmeleitfähigkeit, die entsprechend niedrige Kühlleistungen ermöglichen. In bestimmten Flugkörpern kann aufgrund der hohen Temperaturfestigkeit von C/SiC und C/C und SiC/SiC ganz auf eine Kühlung oder thermische Isolierung verzichtet werden.Particularly noteworthy are the large variety of geometries and shapes at the same time low weight, as well as the excellent temperature resistance and high or adjustable thermal conductivity, which enable correspondingly low cooling capacities. In certain missiles, due to the high temperature resistance from C / SiC and C / C and SiC / SiC entirely to cooling or thermal insulation to be dispensed with.

Man unterscheidet C/SiC und C/C und SiC/SiC mit kontinuierlicher Faserverstärkung sowie kurzfaserverstärktes C/SiC und C/C und SiC/SiC. Das erstgenannte Material aus C/SiC oder C/C oder SiC/SiC, das laminiert, gepreßt oder gewickelt werden kann, zeichnet sich durch besonders hohe Festigkeit und besonders niedrige Dichte aus. Zur Erhöhung der Oxidationsbeständigkeit kann mit einer Oberflächenversiegelung gearbeitet werden. Vorzugsweise werden dafür Schutzschichten aus Siliciumcarbid und/oder Siliciumdioxid und/oder Molybdändisilizid auf die Bauteiloberflächen aufgebracht. Letztere ist bei kurzfaserverstärktem C/SiC überflüssig, da das Material besonders oxidations- und korrosionsbeständig ist. Auch verfügt es über eine extrem gute Wärmeleitfähigkeit und zeichnet sich durch besonders hohe Thermoschockfestigkeit aus. Es eignet sich vor allen Dingen für eine mechanische Bearbeitung im Grünzustand. Dabei können aus C/SiC-Rohlingen und/oder C/C-Rohlingen Bugspitzen 1, feste Flossen 2 oder bewegliche Ruder [Fins] 3, Strahlruder 4, Schubdüsen oder Düsenhalseinsätze 5, Brennkammerauskleidungen 6, Heckkonus 7, Gitterflügel 8, Fluidikelemente 9 und Radome 10 oder Teilkomponenten aus diesen in beliebiger Geometrie aus einem Stück oder aus verschiedenen Einzelsegmenten durch mechanische Bearbeitung leicht geformt werden.A distinction is made between 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. It also has an extreme good thermal conductivity and is characterized by particularly high thermal shock resistance out. It is particularly suitable for mechanical processing in the Green state. 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.

Vorteilhafterweise sind die Einzelsegmente der Bugspitze 1, festen Flossen 2 oder beweglichen Ruder [Fins] 3, Strahlruder 4, Schubdüsen und Düsenhalseinsätze 5, Brennkammerauskleidungen 6, Heckkonus 7, Gitterflügel 8, Fluidikelemente 9 und der Radome 10 oder Teilkomponenten aus diesen mit Siliciumcarbid und/oder Kohlenstoff und/oder Silicium zusammeninfiltriert oder zusammensiliziert, um die gewünschte monolithische Struktur zu geben. Diese Konstruktion eignet sich insbesondere für C/SiC oder C/C oder SiC/SiC mit Kurzfaserverstärkung, wobei die Einzelsegmente vor dem Zusammensilizieren bzw. Infiltrieren mechanisch bearbeitet werden. Eine derartige Flugkörperkomponente 1-10 kann ohne weiteres auch mit Befestigungselementen wie z.B. Schrauben oder Bolzen oder Flanschen, vorzugsweise aus C/SiC und/oder C/C und/oder SiC/SiC, verbunden werden. Außerdem können in die Flugkörperkomponenten 1-10 durch mechanische Bearbeitung im Grünzustand Kühlkanäle und/oder Ausnehmungen mit runden, rechteckigen oder schlitzförmigen Querschnitt eingebracht werden.Advantageously, the individual segments of the bow tip 1, fixed fins 2 or movable rudder [fins] 3, thruster 4, thrusters and nozzle neck inserts 5, Combustion chamber linings 6, rear cone 7, grille wing 8, fluidic elements 9 and of the radomes 10 or partial components made of these with silicon carbide and / or carbon and / or silicon infiltrated or collated to the desired one to give monolithic structure. 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. In addition, in the missile components 1-10 by mechanical processing in the green state Cooling channels and / or recesses with round, rectangular or slit-shaped Cross section are introduced.

Das erfindungsgemäße Verfahren sieht eine Gestaltung der Flugkörperkomponenten 1-10 in Hybrid- und Segmentbauweise vor. Durch mechanische Bearbeitung von Rohkörpern und Teilsegmenten, die vorzugsweise aus C/SiC und/oder C/C und/oder SiC/SiC oder aus geeigneten Kombinationen mit kontinuierlicher Faserverstärkung und/oder Kurzfaserverstärkung besteht und durch die anschließende Infiltration mit Silicium und/oder Siliciumcarbid und/oder Kohlenstoff dieser Einzelsegmente werden monolithische Flugkörperkomponenten in Hybridbauweise ausgebildet.The method according to the invention provides for a design of the missile components 1-10 in hybrid and segment design. By mechanical processing of 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.

Die Innenwand der Flugkörper oder die thermisch hochbelasteten Stellen der Flugkörper kann in geeigneter Weise mit C/SiC- oder C/C- oder SiC/SiC-Segmenten ausgekleidet werden und mittels Kühlung über Kühlkanäle und/oder mit einem Isolationsmaterial, vorzugsweise aus C/SiC oder C/C oder SiC/SiC oder aus Kohlenstoffaserfilzen oder Graphitfolie oder Kombinationen aus diesen, das die Temperatur- und Druckbelastung der metallischen Flugkörperstruktur soweit wie möglich reduziert, versehen sein und zu einer monolithischen Flugkörperkomponente 1-10 zusammensiliziert werden. Die Isolationswerkstoffe können auch unter Zwischenschaltung von Abstandshaltern, vorzugsweise aus C/SiC oder C/C oder SiC/SiC, mit den Flugkörperkomponenten 1-10 aus C/SiC und/oder C/C miteinander verbunden werden, um die gewünschte monolithische Struktur zu ergeben.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.

Vorteilhafterweise kann die Dichte und Porosität des C/SiC- und/oder des C/C- und/oder SiC/SiC-Materials während der Infiltration oder Silizierung durch die Zugabemenge an Silicium, Kohlenstoff oder Siliciumcarbid eingestellt werden, sodaß das C/SiC und/oder C/C und/oder SiC/SiC mit hoher Dichte und geringer Porosität als thermomechanische Tragstruktur und/oder Auskleidung und das C/SiC und/oder C/C und/oder SiC/SiC mit niedriger Dichte bzw. hoher Porosität als Wärmeisolierung eingesetzt werden kann. Dabei können auch Dichte- und Porositätsgradienten über der Wandstärke der Flugkörperkomponenten 1-10 eingestellt werden. Advantageously, 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.

Durch die Gas- und Flüssigkeitsdichtigkeit der C/SiC- und/oder C/C-Materialien können in die metallische Flugkörperstruktur auch offene Kühlkanäle eingearbeitet werden, die beim Einsetzen der C/SiC- und/oder C/C-Teile und/oder SiC/SiC-Teile geschlossen werden. Die Flugkörperkomponente 1-10 wird je nach verwendetem System aus C/SiC- und/oder C/C- und/oder SiC/SiC-Einzelsegmenten gefertigt, die anschließend zu einer monolithischen Struktur mit Kohlenstoff und/oder Silicium und/oder Siliciumcarbid zusammeninfiltriert und/oder zusammensiliziert werden oder man fertigt die Flugkörperkomponenten 1-10 aus einem Stück, vorzugsweise durch mechanische Bearbeitung eines C/SiC- und/oder C/C- und/oder SiC/SiC-Rohlings. Diese C/SiC- und/oder C/C-Teile und/oder SiC/SiC-Teile können auch die Kühlkanäle (falls notwendig) oder Ausnehmungen bereitstellen, um die Wärme abzutransportieren. Der C/SiC- und/oder C/C-Körper und/oder SiC/SiC-Körper 1-10 und die metallische Flugkörperstruktur sind mit geeigneten Verbindungselementen wie z.B. Bolzen-, Schraub- oder Flanschverbindungen, vorzugsweise aus C/SiC und/oder C/C und/oder SiC/SiC, miteinander zu verbinden. Möglichkeiten hierzu sind in den Bildern 2 bis 9 gezeigt.Due to the gas and liquid tightness of the C / SiC and / or C / C materials open cooling channels are also incorporated into the metallic missile structure, closed when inserting the C / SiC and / or C / C parts and / or SiC / SiC parts become. 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. These 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.

Die Erfindung wird im folgenden, anhand bevorzugter Ausführungsbeispiele im Zusammenhang mit den beiliegenden Zeichnungen, näher erläutert. Durch den Einsatz von Flugkörperkomponenten 1-10 aus faserverstärkter Keramik (C/SiC und/oder C/C und/oder SiC/SiC) kommt es zu einer erheblichen Gewichtsreduzierung im Vergleich zu metallischen Flugkörperkomponenten. Durch die Hochtemperaturfestigkeit von C/SiC und/oder C/C und/oder SiC/SiC kann auf die Kühlung ganz oder teilweise verzichtet werden. Das erfindungsgemäße Verfahren erlaubt jegliche Geometrie- und Größenvariationen bei den Flugkörperkomponenten 1-10. The invention is in the following, based on preferred embodiments in connection with the accompanying drawings, explained in more detail. Because of the engagement of missile components 1-10 made of fiber-reinforced ceramic (C / SiC and / or C / C and / or SiC / SiC) there is a considerable reduction in weight in comparison to metallic missile components. Due to the high temperature resistance of C / SiC and / or C / C and / or SiC / SiC can affect the cooling in whole or in part to be dispensed with. The method according to the invention allows any geometry and size variations for missile components 1-10.

In Bild 1 ist ein Flugkörper nach derzeitigem Stand der Technik dargestellt. Aufgrund der hohen Temperatur- und Druckbelastung kommen derzeit z.B. nur warmfeste Metalle und Metallegierungen mit hoher Dichte infrage, die aufgrund ihrer relativ geringen Temperaturfestigkeit gekühlt werden müssen. Neben diesen thermomechanischen Anforderungen müssen die metallischen Werkstoffe auch allen Anforderungen bezüglich Korrosion, Bearbeitung, Oberflächengüte und Schweißbarkeit genügen. 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.

In Bild 2 ist eine Bugspitze 1 und ein Radom 10 eines Flugkörpers dargestellt. Die Bugspitze wird besonders durch hohe Drücke und hohe Temperaturen beansprucht. Durch den Einsatz von faserverstärkter Keramik kann das Gewicht der Bugspitze um mindestens 1 kg im Vergleich zu einer metallischen Bugspitze reduziert werden.
An Radomen treten Beanspruchungen durch hohe Drücke und hohe Temperaturen auf. Zusätzlich ist bei Radomen eine erhöhte Radardurchlässigkeit und Oberflächengenauigkeit (z.B: durch Schleifbarkeit) sowie der Aufbau unterschiedlicher Wandstärken erforderlich. 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. By using fiber-reinforced ceramics, 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. In addition, with radomes, increased radar permeability and surface accuracy (e.g. through grindability) and the construction of different wall thicknesses are required.

In Bild 3 sind die Stabilisierungsfins bzw. festen Flossen 2 und der Heckkonus 7 eines Flugkörpers dargestellt.
An den festen Flossen treten vor allem Beanspruchungen durch hohe Längs- und Querbeschleunigungskräfte und durch hohe Temperaturen auf. Der Heckkonus 7 eines Flugkörpers wird durch hohe Drücke und hohe Temperaturen beansprucht und dient zur Stabilisierung des Flugkörpers. Der Einsatz von faserverstärkter Keramik führt am Heckkonus zu einer Gewichtsersparnis von 3 kg. 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.

In Bild 4 sind bewegliche Ruder bzw. Fins 3 und Gitterflügel 8 dargestellt. An den beweglichen Rudern bzw. Fins 3 treten Beanspruchungen durch hohe Längs- und Querbeschleunigungskräfte und durch hohe Temperaturen auf. Sie dienen als aerodynamische Lenkhilfe. Auch an den Gitterflügeln 8 treten Beanspruchungen durch hohe Längs- und Querbeschleunigungskräfte und durch hohe Temperaturen auf. Sie dienen sowohl als aerodynamische Lenkhilfe als auch zur Erhaltung der Stabilität des Flugkörpers. Der Gitterflügel sieht aus wie ein am Heck des Flugkörpers angebrachter schmaler Fußabstreifer, dessen Öffnungen in Flugrichtung stehen und um die Längsachse gedreht werden kann. 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.

In Bild 5 sind die erfindungsgemäßen Strahlruder 4 abgebildet. Eine Beanspruchung durch hohe Querkräfte, Temperaturen und Abrasion durch Abgase und Feststoffteilchen (z.B. Al2O3-Partikeln) muß bei der Auslegung von Strahlrudern berücksichtigt werden. Der Einsatz von Strahlrudern im Abgasstrahl dient als zusätzliche Lenkhilfe während der Antriebsphase des Flugkörpers. Strahlruder, die im hinteren Bereich einer Raketendüse direkt im Abgasstrahl zur Strahlumlenkung eingebaut sind, unterliegen extrem hohen thermo-mechanischen Beanspruchungen durch die heißen, reaktiven Verbrennungsgase und den hohen Querkräften. Thermoschockbeständigkeit und ein gutes Abrasionsverhalten gegenüber Feststoffteilchen, wie z.B. Al2O3 und Ruß, werden bei Strahlrudern zusätzlich gefordert, da Strahlruder je nach Motortyp und Treibstoffart Gas/Teilchenströmungen mit Temperaturen von 2500 °C plötzlich ausgesetzt sein können. In 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, such as Al 2 O 3 and soot, 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.

In Bild 6 sind eine Schubdüse 5 und die typische Ausführungsform der erfindungsgemäßen Brennkammerauskleidung 6 dargestellt. Figure 6 shows a thrust nozzle 5 and the typical embodiment of the combustion chamber lining 6 according to the invention.

Die Schubdüse wird durch extrem hohe Drücke und Temperaturen beansprucht. Oft besitzen die Triebwerke von Flugkörpern für die einzelnen Schubphasen (Auswurf-, Beschleunigungs- und Marschphase) mehrere und unterschiedlich viele Schubdüsen.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.

In Bild 7 sind typische Fluidikelemente 9 abgebildet, die als Querschubsteuerungen eingesetzt werden.
Das erfindungsgemäße Verfahren sieht vor, daß man die Schubdüse und/oder den Düsenhals und/oder die Brennkammer mit C/SiC- und/oder C/C-Segmenten und/oder SiC/SiC-Segmenten auskleidet. Die Innenwände der Flugkörper sind aus C/SiC- und/oder C/C- und/oder SiC/SiC-Einzelsegmenten gestaltet. Die C/SiC- und/oder C/C- und/oder SiC/SiC-Segmente sind so zu gestalten, daß die Teilungsschlitze, die unter hohem Druck und hoher Temperatur stehenden Gase nicht zur metallischen Flugkörperstruktur durchgelassen werden. Die C/SiC- und/oder C/C-Teile und/oder SiC/SiC-Teile können der Innenkontur des Flugkörpermotors angepaßt werden und ermöglichen so eine geometrische Vereinfachung der Flugkörperstruktur. 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.

Das Verfahren sieht vor, daß man die C/SiC- und/oder C/C- und/oder SiC/SiC-Einzelsegmente für die Flugkörperkomponenten (1-10) aus C/SiC- und/oder C/C- und/oder SiC/SiC-Rohlingen mechanisch bearbeitet und vor der Montage in die Flugkörperstruktur zu einer monolithischen Struktur zusammensiliziert.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.

In den Beispielen kann die Kühlung (falls notwendig) wahlweise über das Einbringen von Kühlkanälen oder Ausnehmungen in die C/SiC- und/oder C/C- und/oder SiC/SiC-Struktur oder die Isolation mit Kohlenstoffilzen oder Graphitfolie oder C/SiC oder C/C oder SiC/SiC oder Kombinationen aus diesen erfolgen. Die Kühlung mit Kühlkanälen kann wahlweise je nach Anforderung in der Flugkörperstruktur am Übergang Metall zu C/SiC und/oder C/C und/oder SiC/SiC oder im C/SiC- und/oder C/C- und/oder SiC/SiC-Teil selbst erfolgen. Es ist auch eine Kombination aus beiden Teilen vorgesehen.In the examples, the cooling (if necessary) 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.

Claims (3)

  1. Process for manufacturing missiles or missile components subjected to high thermal and mechanical loading, with the following steps:
    a) producing blanks from fibre-reinforced green ceramic, namely carbon-fibre-reinforced silicon carbide (C/SiC) and/or carbon-fibre-reinforced carbon (C/C) and/or silicon-carbide-fibre-reinforced silicon carbide (SiC/SiC);
    b) shaping and machining of the ceramic blanks in the green state to correspond to the geometry of the parts of the missile or the missile components;
    c) joint infiltration or joint siliconization of the blanks with silicon and/or silicon carbide and/or carbon to obtain a monolithic composite structure, the joined-together blanks forming the overall form of the missile or of the respective missile component.
  2. Process according to Claim 1, characterized in that the density and porosity of the S/SiC [sic] and/or the C/C and/or SiC/SiC material is set during the infiltration or siliconization step by adding an amount of silicon, carbon and/or silicon carbide, so that there is obtained a supporting structure with on the one hand high density and low porosity as a supporting part and/or lining and with on the other hand low density and high porosity as thermal insulation.
  3. Process according to Claim 1, characterized in that cooling channels are made or formed into the ceramic blanks by machining.
EP97935567A 1996-08-16 1997-08-04 Process for manufacturing missiles or missile components Expired - Lifetime EP0918976B1 (en)

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DE19632893 1996-08-16
DE19632893A DE19632893C2 (en) 1996-08-16 1996-08-16 Process for manufacturing missile components from fiber-reinforced ceramic
PCT/EP1997/004235 WO1998008044A1 (en) 1996-08-16 1997-08-04 Missile components made of fibre-reinforced ceramics

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EP0918976B1 true EP0918976B1 (en) 2000-06-14

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US7037602B2 (en) 2002-07-04 2006-05-02 Sgl Carbon Ag Multilayer composite
DE10230231B4 (en) * 2002-07-04 2007-07-05 Sgl Carbon Ag Multilayer composite material

Also Published As

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

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