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1. 发明授权

US12065991B2 High temperature thermal protection system for rockets, and associated methods 有权
公开(公告)号：US12065991B2
公开(公告)日：2024-08-20
申请号：US17514471
申请日：2021-10-29
申请人： Blue Origin, LLC
发明人： Adam Keith Norman , John Paul Brendel , Christopher Patrick Hupf , Stefano Gulli
IPC分类号： B64G1/58 , B32B5/02 , B32B5/26 , B32B7/027 , B32B7/09 , B32B7/12 , B64G1/22 , F02K9/97
CPC分类号： F02K9/974 , B32B5/024 , B32B5/26 , B32B7/027 , B32B7/09 , B32B7/12 , B64G1/226 , B64G1/58 , F02K9/97 , F02K9/972 , B32B2262/103 , B32B2262/105 , B32B2307/304 , B32B2605/00 , B32B2605/18
摘要： A high temperature thermal protection systems for rockets, and associated methods, is disclosed. A representative system includes a launch vehicle having a first end and a second end generally opposite the first end. The launch vehicle is elongated along a vehicle axis extending between the first and second ends and carries a propulsion system having at least one nozzle positioned at the second end of the launch vehicle. A thermal protection apparatus positioned around the nozzle is used to provide cooling and/or insulation to the nozzle during the flight of the launch vehicle. The thermal protection apparatus can include multiple fabric layers and an insulation layer stacked and stitched together. The fabric layers can include metal alloy fibers. In representative systems, the thermal protection apparatus can further include provisions for water that saturates the insulation layer to provide further insulating and/or cooling effects.

2. 发明授权

US12031507B2 Augmented aerospike nozzle, engine including the augmented aerospike nozzle, and vehicle including the engine 有权
公开(公告)号：US12031507B2
公开(公告)日：2024-07-09
申请号：US17754787
申请日：2020-08-27
申请人： STOKE Space Technologies, Inc.
发明人： Andrew Lapsa , Thomas Feldman
IPC分类号： F02K9/97 , B64G1/00 , F02K9/64
CPC分类号： F02K9/97 , B64G1/006 , F02K9/64 , F02K9/972 , F05D2240/1281
摘要： An augmented aerospike nozzle includes a throat, a centerbody extending aft of the throat, an inner expansion surface defined by the centerbody, an outer expansion surface outboard of the inner expansion surface, and an expansion cavity defined between the inner expansion surface and the outer expansion surface. An engine includes a high pressure chamber and the augmented aerospike nozzle. A vehicle for supersonic flight includes the engine with the augmented aerospike nozzle.

3. 发明授权

US11976613B1 Dual mode rocket engine with bipropellant injection in the nozzle 有权
公开(公告)号：US11976613B1
公开(公告)日：2024-05-07
申请号：US18457283
申请日：2023-08-28
申请人： Pivotal Space, Inc.
发明人： Lloyd J. Droppers
IPC分类号： F02K9/82 , F02K9/44 , F02K9/56 , F02K9/80 , F02K9/64 , F02K9/97
CPC分类号： F02K9/82 , F02K9/44 , F02K9/56 , F02K9/80 , F02K9/64 , F02K9/972
摘要： A rocket propulsion system that may include a supersonic rocket nozzle with a supersonic divergent section, and a heat transfer system configured to transfer heat from the supersonic rocket nozzle to a propellant where a portion of the propellant may be selectively injected, combusted, and expanded in the supersonic nozzle generating an additional thrust. In examples, the heated propellant may be used to power a pump system to feed the rocket engine.

4. 发明公开

US20230332562A1 STAGED COMBUSTION LIQUID ROCKET ENGINE CYCLE WITH THE TURBOPUMP UNIT AND PREBURNER INTEGRATED INTO THE STRUCTURE OF THE COMBUSTION CHAMBER 审中-公开
公开(公告)号：US20230332562A1
公开(公告)日：2023-10-19
申请号：US18211277
申请日：2023-06-18
申请人： Special Aerospace Services, LLC
发明人： Timothy Bulk , Christopher Hayes
IPC分类号： F02K9/48 , F02K9/52 , F02K9/82 , F02K9/64
CPC分类号： F02K9/48 , F02K9/52 , F02K9/82 , F02K9/64 , F02K9/972
摘要： Devices and methods of rocket propulsion are disclosed. In one aspect, a staged combustion liquid rocket engine with preburner and turbopump unit (TPU) integrated into the structure of the combustion chamber is described. An initial propellant mixture is combusted in a preburner combustion chamber formed as an annulus around a main combustion chamber, the combustion products from the preburner driving the turbine of the TPU and subsequently injected into the main combustion chamber for secondary combustion along with additional propellants, generating thrust through a supersonic nozzle. The preburner inner cylindrical wall is shared with the outer cylindrical wall of the engine's main combustion chamber and the turbine is axially aligned with the main combustion chamber. Liquid propellants supplied to the engine are utilized for regenerative cooling of the combustion chamber and preburner, where the liquid propellants are gasified in cooling manifolds before injection into the preburner and main combustion chamber.

5. 发明授权

US10495028B1 Thermoelectric rocket propellant tank pressurization system 有权
公开(公告)号：US10495028B1
公开(公告)日：2019-12-03
申请号：US16209213
申请日：2018-12-04
申请人： Vector Launch Inc.
发明人： Mark Christopher Sipperley , Colin Healey Smith
IPC分类号： F02K9/56 , F02K9/46 , F02K9/50 , F02K9/64 , F02K9/60 , F02K9/97
CPC分类号： F02K9/563 , F02K9/46 , F02K9/50 , F02K9/605 , F02K9/64 , F02K9/972 , F05D2270/303
摘要： Various enhanced rocket engine systems are discussed herein. In one implementation, a rocket engine system includes a combustion chamber, and at least one propellant tank that holds propellant in at least a liquid state. The rocket engine system also includes a pump configured to pump liquid propellant from the at least one propellant tank through a thermoelectric generator (TEG) system and a heat exchanger. The TEG system is configured to produce electrical power for the pump based at least on a temperature differential between the liquid propellant from the at least one propellant tank and heat produced in the combustion chamber during an active state of the rocket engine. The heat exchanger is configured to receive heat from the combustion chamber and pressurize the at least one propellant tank by heating at least partially liquid propellant received from the TEG system.

6. 发明授权

US09291410B2 Cooling device for cooling combustion gases from recoilless anti-tank weapons 有权
标题翻译：用于冷却反坦克武器燃烧气体的冷却装置
公开(公告)号：US09291410B2
公开(公告)日：2016-03-22
申请号：US13811036
申请日：2010-07-22
申请人： Lars Nilsson
发明人： Lars Nilsson
IPC分类号： F02K11/00 , F41A1/10 , F02K9/40 , F02K9/97 , F41A13/00 , F41A13/04
CPC分类号： F41A1/10 , F02K9/40 , F02K9/972 , F41A13/00 , F41A13/04
摘要： The present invention relates to a cooling device (2) for cooling combustion gases from a rocket motor (1) in an antitank weapon. The cooling device (2) comprises a coolant reservoir (3, 21) containing a coolant (4). The coolant reservoir (3, 21) is arranged such that the coolant (4) is transferred from the coolant reservoir (3, 21) to the combustion gases in the gas outlets (6) of the rocket motor (1) in response to a pressure increase in the rocket motor (1). The coolant reservoir (3, 21) constitutes an integral part of the rocket motor (1) and is connected to the combustion chamber (7) by at least two gas inlets (8) for pressurization of the coolant (4). The coolant reservoir is furthermore connected to the rocket motor nozzle (9) by at least two coolant outlets (10) for transfer of coolant (4) from the coolant reservoir (3, 21) to the rocket motor nozzle (9).
摘要翻译：本发明涉及一种用于冷却来自反坦克武器的火箭发动机（1）的燃烧气体的冷却装置（2）。冷却装置（2）包括容纳冷却剂（4）的冷却剂储存器（3,21）。冷却剂储存器（3,21）被布置成使得冷却剂（4）响应于冷却剂储存器（3,21）从冷却剂储存器（3,21）转移到火箭发动机（1）的气体出口（6）中的燃烧气体火箭发动机（1）的压力增加。冷却剂储存器（3,21）构成火箭发动机（1）的一体部分，并且通过至少两个用于冷却剂（4）加压的气体入口（8）连接到燃烧室（7）。冷却剂储存器还通过至少两个用于将冷却剂（4）从冷却剂储存器（3,21）传递到火箭发动机喷嘴（9）的冷却剂出口（10）连接到火箭发动机喷嘴（9）。

7. 发明授权

US09222439B2 Rocket engine with cryogenic propellants 有权
标题翻译：火箭发动机带低温推进剂
公开(公告)号：US09222439B2
公开(公告)日：2015-12-29
申请号：US13384477
申请日：2010-07-16
申请人： Olivier Dobek , Daniel Le Dortz
发明人： Olivier Dobek , Daniel Le Dortz
IPC分类号： F02K9/50 , F02K9/97 , F02K9/64
CPC分类号： F02K9/972 , F02K9/50 , F02K9/64
摘要： A cryogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid.
摘要翻译：低温推进剂火箭发动机包括：用于第一液体推进剂的至少第一罐; 用于第二液体推进剂的第二罐; 用于惰性流体的第三罐; 包括燃烧室的轴对称喷嘴，用于将第一和第二液体推进剂喷射到燃烧室中的装置，喷嘴喉部和发散部分; 以及加热器装置，其包括用于输送惰性流体并且布置在喷嘴外部的至少一个管道，其直接邻近但不与其接触，以回收当火箭发动机运行时发射的热辐射的能量并且加热惰性流体。

8. 发明授权

US09127622B2 Reversible flow discharge orifice 有权
标题翻译：可逆流量排放口
公开(公告)号：US09127622B2
公开(公告)日：2015-09-08
申请号：US13300775
申请日：2011-11-21
申请人： Jim A. Clark , Craig W. Irwin , Reed A. Kakuska
发明人： Jim A. Clark , Craig W. Irwin , Reed A. Kakuska
IPC分类号： F02K9/56 , F16L55/027
CPC分类号： F02K9/566 , F02K9/64 , F02K9/972 , F15D1/025 , F16L55/027 , G05D7/0186
摘要： A rocket engine fluid-flow system includes a pump fluidly interconnecting a fluid source to a combustion chamber. A nozzle is in fluid communication with the combustion chamber and includes coolant tubes fluidly arranged between the pump and the combustion chamber. An orifice has a throat and is fluidly arranged between the pump and the coolant tubes. The orifice has entrance and exit ramps arranged on either side of the throat. The exit ramp has an exit ramp surface with a divergent angle that is less than a right angle. The entrance ramp provides a smooth approach to the orifice throat. In one example, the exit ramp includes an exit ramp surface having a divergent angle of 20-60°. The exit ramp radius is less than twice the throat radius in one example.
摘要翻译：火箭发动机流体流动系统包括将流体源流体地连接到燃烧室的泵。喷嘴与燃烧室流体连通，并且包括流体地布置在泵和燃烧室之间的冷却剂管。孔具有喉部并且流体地布置在泵和冷却剂管之间。孔口具有设置在喉部两侧的入口和出口斜面。出口斜坡具有出口斜面，其具有小于直角的发散角。入口斜坡提供了孔口喉咙平滑的方法。在一个示例中，出口斜坡包括具有20-60°的发散角的出口斜面。在一个示例中，出口斜坡半径小于喉部半径的两倍。

9. 发明授权

US09115666B2 Slotted multi-nozzle grid with integrated cooling channels 有权
标题翻译：带有集成冷却通道的开槽多喷嘴格栅
公开(公告)号：US09115666B2
公开(公告)日：2015-08-25
申请号：US13647599
申请日：2012-10-09
申请人： Raytheon Company
发明人： Daniel Chasman , Stephen D. Haight , James K. Villarreal
IPC分类号： F02K9/30 , F02K9/97
CPC分类号： F02K9/30 , F02K9/97 , F02K9/972 , Y10T29/49346
摘要： An apparatus includes a slotted multi-nozzle grid with a plate having multiple elongated slotlettes through the plate. Each of at least some of the slotlettes has a convergent input, a divergent output, and a narrower throat portion separating the convergent input and the divergent output. At least some of the slotlettes are arranged in multiple rows. The plate further includes multiple cooling channels through the plate. At least some of the cooling channels are located between the rows of slotlettes. Each cooling channel is configured to transport coolant through the plate in order to cool the plate, such as to cool the plate as hot combustion gases pass through the plate. Each of at least some of the rows may include at least two slotlettes, and two adjacent slotlettes in one row may be separated by a structural ligament (which may have a tear-drop cross-sectional shape).
摘要翻译：一种装置包括带槽的多喷嘴格栅，其具有穿过该板的具有多个细长槽的板。至少一些小缝隙中的每一个具有收敛输入，发散输出和分离收敛输入和发散输出的较窄喉部。至少一些小行星排列成多行。板还包括穿过板的多个冷却通道。至少一些冷却通道位于行槽之间。每个冷却通道被构造成运输冷却剂通过板以便冷却板，例如当热燃气通过板时冷却板。至少一些行中的每一行可以包括至少两个开槽元件，并且一行中的两个相邻的开槽元件可以被结构韧带（其可能具有泪滴横截面形状）分开。

10. 发明授权

US08708647B2 Liner for a turbine section, a turbine section, a gas turbine engine and an aeroplane provided therewith 有权
标题翻译：用于涡轮机部分的涡轮机部分，涡轮机部分，燃气涡轮发动机和与其配备的飞机
公开(公告)号：US08708647B2
公开(公告)日：2014-04-29
申请号：US12518101
申请日：2006-12-06
申请人： Arne Boman
发明人： Arne Boman
IPC分类号： F01D25/12
CPC分类号： F02K9/64 , F02C7/12 , F02K1/82 , F02K1/822 , F02K1/825 , F02K9/972
摘要： A liner for a turbine section includes a first wall, a plurality of webs interconnected with and projecting from the first wall, and a plurality of cooling channels, each of the cooling channels being delimited by two adjacent webs and the first wall, wherein each cooling channel presents a height corresponding to the height of its delimiting webs, and a width corresponding to the distance between its delimiting webs. At least one of the cooling channels has a width/height ratio of below 5 or/and the material of the webs has a higher thermal conductivity than the material of the first wall. A turbine section, a gas turbine engine and an aeroplane provided with such a liner are also disclosed.
摘要翻译：用于涡轮部分的衬套包括第一壁，与第一壁互连并且从第一壁突出的多个腹板和多个冷却通道，每个冷却通道由两个相邻腹板和第一壁限定，其中每个冷却通道具有对应于其分隔腹板的高度的高度，以及对应于其分隔腹板之间的距离的宽度。冷却通道中的至少一个具有低于5的宽度/高度比或/和幅材的材料具有比第一壁的材料更高的热导率。还公开了一种涡轮机部分，燃气涡轮发动机和设有这种衬管的飞机。

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