Rocket engine
IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0476568
(2003-02-27)
|
우선권정보 |
FR-0002686 (2002-03-04) |
국제출원번호 |
PCT//FR03/00630
(2003-11-04)
|
§371/§102 date |
20031104
(20031104)
|
국제공개번호 |
WO03//07485
(2003-09-12)
|
발명자
/ 주소 |
|
출원인 / 주소 |
- Eads Space Transportation SA
- MBDA France
|
대리인 / 주소 |
Stevens, Davis, Miller &
|
인용정보 |
피인용 횟수 :
2 인용 특허 :
9 |
초록
▼
The invention concerns a rocket engine wherein the combustion chamber includes at least one first monolithic component made of a thermostructural composite material comprising a porous wall through which the fuel is introduced in the core of the combustion chamber. A small part of the fuel is direct
The invention concerns a rocket engine wherein the combustion chamber includes at least one first monolithic component made of a thermostructural composite material comprising a porous wall through which the fuel is introduced in the core of the combustion chamber. A small part of the fuel is directed towards the neck for it to be cooled.
대표청구항
▼
1. A rocket engine comprising a combustion chamber in the heart of which a fuel and an oxidizer are burnt and which is connected, by a throat, to a divergent nozzle through which the gases resulting from said combustion escape, said heart being supplied with oxidizer via its opposite end to said thr
1. A rocket engine comprising a combustion chamber in the heart of which a fuel and an oxidizer are burnt and which is connected, by a throat, to a divergent nozzle through which the gases resulting from said combustion escape, said heart being supplied with oxidizer via its opposite end to said throat and being surrounded by a porous skin of thermostructural composite which receives fuel on its opposite outer side to said heart, some of this fuel being introduced into said heart through said porous skin, wherein said proportion of the fuel introduced into said heart through said porous skin constitutes the fuel supply to said engine and in that the proportion of said fuel not passing through said porous skin is directed toward said throat to cool it.2. The rocket engine as claimed in claim 1, wherein said porous skin forms part of a first monolithic piece of thermostructural composite comprising two skins of composite spaced apart from one another leaving between them an intermediate space and joined together by a plurality of threadlike spacers of composite.3. The rocket engine as claimed in claim 2, which is provided with a longitudinal axis and in which said divergent nozzle is arranged in the continuation of said combustion chamber, on the opposite side of said throat to said combustion chamber, wherein:said first monolithic piece is cylindrical and arranged coaxially with respect to said longitudinal axis so that one of said skins is an inner skin whereas the other is an outer skin; said oxidizer is introduced into the cylindrical volume delimited by said inner skin and forming the heart of said combustion chamber, on the opposite side to said nozzle; and said fuel is introduced into said intermediate space, which therefore has an annular cross section, also on the opposite side to said nozzle. 4. The rocket engine as claimed in claim 3, wherein said outer skin of said first monolithic piece is sealed against liquids and against gases.5. The rocket engine as claimed in claim 3, wherein said first monolithic piece has an inside diameter greater than that of said throat and in that the annular orifice of said intermediate space, arranged on the same side as said nozzle, lies facing the convergent part of said throat.6. The rocket engine as claimed in claim 2, wherein said nozzle comprises, beyond said throat, a sheath able to house said first monolithic piece.7. The rocket engine as claimed in claim 2, wherein said nozzle consists of a second monolithic piece of composite.8. The rocket engine as claimed in claim 2, wherein said nozzle consists of a second monolithic piece of composite and in that said second monolithic piece constitutes a continuation of said outer skin of said first monolithic piece, this continuation forming an integral part of said outer skin.9. The rocket engine as claimed in claim 2, which is provided with a longitudinal axis and in which said combustion chamber is arranged in said divergent nozzle near the vertex thereof, wherein:said combustion chamber comprises: an inner first monolithic piece of composite, of cylindrical shape, arranged coaxially with respect to said axis and having an inner composite skin and an outer composite skin; and an outer first monolithic piece of composite, of cylindrical shape, arranged coaxially with respect to said axis and having an inner composite skin and an outer composite skin, said outer first piece surrounding said inner first piece, so as to form between them an annular heart for said combustion chamber; said inner and outer first pieces forming between them and the vertex of said divergent nozzle an annular passage for communication with said nozzle; said oxidizer is introduced into the annular heart of said combustion chamber from the opposite side to said vertex of the nozzle; and said fuel is introduced into the intermediate spaces, of annular cross section, of said inner and outer first pieces from the opposite side to said vertex of the nozzle. 10. The rocket engine as claimed in claim 9, wherein the inner skin of the inner first piece is sealed against liquids and against gases.11. The rocket engine as claimed in claim 9, wherein the vertex of said divergent nozzle is pierced with an orifice and in that the collection of said inner and outer first pieces is secured to said nozzle by a third monolithic piece of composite in the shape of a horn.12. The rocket engine as claimed in claim 9, wherein said combustion chamber is supplied with fuel via a dome-shaped piece arranged on the opposite side of said combustion chamber to the vertex of the nozzle and the convex wall of which faces toward the nozzle and is made of thermostructural composite.
이 특허에 인용된 특허 (9)
-
Vuillamy Didier (Quincampoix FRX) Tiret Etienne (La Chapelle Reanville FRX) Beurain Andr (Chambly FRX), Close combustion gas generator.
-
Haidn Oskar,DEX ; Hald Hermann,DEX ; Lezuo Michael,DEX ; Winkelmann Peter,DEX, Combustion chamber for rocket engine.
-
Schmidt Guenther,DEX ; Beyer Steffen,DEX, Combustion chamber wall construction for high power engines and thrust nozzles.
-
Joachim Kretschmer DE, Coolable nozzle and method for producing such a nozzle for a rocket engine.
-
Pellet Marc (Vernon FRX), Enclosure containing hot gases cooled by transpiration, in particular the thrust chamber of a rocket engine.
-
Filipuzzi Ludovic (Pessac FRX) Huet Philippe (Margaux FRX), Method for making a sealed passage in a refractory composite part, and application to the production of a refractory com.
-
Coffinberry George A. (West Chester OH), Multiple-propellant air vehicle and propulsion system.
-
Niino Masayuki (Sendi JPX) Yatsuyanagi Nobuyuki (Shibata JPX) Kumakawa Akiraga (Souma JPX) Suzuki Akio (Shibata JPX) Gomi Hiromi (Shibata JPX) Sakamoto Hiroshi (Shibata JPX) Sasaki Masaki (Shibata JP, Rocket combustion chamber cooling wall of composite cooling type and method of manufacturing the same.
-
Beyer Steffen,DEX ; Wiedmann Dietmar,DEX, Wall construction for a combustion chamber or a nozzle of a high performance propulsion plant.
이 특허를 인용한 특허 (2)
-
Conrardy, Jean Marie, Propulsion assembly and method.
-
Peyrisse, Daniel; Conrardy, Jean-Marie, Thruster comprising a plurality of rocket motors.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.