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公开(公告)号：US12031507B2
公开(公告)日：2024-07-09
申请号：US17754787
申请日：2020-08-27
Applicant: STOKE Space Technologies, Inc.
Inventor： Andrew Lapsa , Thomas Feldman
CPC classification number: F02K9/97 , B64G1/006 , F02K9/64 , F02K9/972 , F05D2240/1281
Abstract: 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.
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公开(公告)号：US11976613B1
公开(公告)日：2024-05-07
申请号：US18457283
申请日：2023-08-28
Applicant: Pivotal Space, Inc.
Inventor： Lloyd J. Droppers
Abstract: 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.
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公开(公告)号：US20240127972A1
公开(公告)日：2024-04-18
申请号：US18257655
申请日：2021-12-20
Applicant: Tokamak Energy Ltd
Inventor： Rob Bamber
Abstract: A plasma-facing component for a plasma chamber, comprising: a plasma-facing target surface; an inlet through which to receive a coolant fluid and an outlet through which to expel the coolant fluid; and a plurality of internal cooling channels. Each cooling channel is connected to the inlet by a plurality of feed channels and to the outlet by a plurality of return channels, the feed channels being configured to direct coolant fluid against a region of a wall of the cooling channel. Respective openings of the feed and return channels into the cooling channel are arranged in non-overlapping repeating units along a length of the cooling channel Each unit comprises openings of at least one feed channel and at least one return channel.
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公开(公告)号：US11846251B1
公开(公告)日：2023-12-19
申请号：US16857609
申请日：2020-04-24
Applicant: FireFly Aerospace Inc.
Inventor： Anatoli Alimpievich Borissov , Thomas Edward Markusic
CPC classification number: F02K7/18 , F02K9/48 , F02K9/64 , F02K9/74 , F05D2220/80 , F05D2240/35 , F05D2260/205
Abstract: The inventors introduce the Borissov-Markusic Cycle as the new rocket engine cycle to solve the problem of low efficient open gas generator or tap-off gas generator cycles used to supply power to turbopump. A liquid rocket engine directs turbopump exhaust from a turbopump to a booster engine having an intake to accept ambient airflow, such as a variation of a ramjet, scramjet or dual mode ram scramjet engine. The turbopump is powered by combustion gases, such as from a gas generator or a tap-off manifold interfaced with the liquid rocket engine combustion chamber, and applies energy of the combustion gases to pump fuel and/or liquid oxygen to the liquid rocket engine combustion chamber. The combustion gases have a fuel-rich composition that includes unconsumed fuel from incomplete oxidation so that, upon injection into the combustion chamber of the booster engine, oxidation by ambient air of the unconsumed fuel releases energy to generate thrust with the booster engine.
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公开(公告)号：US20230332562A1
公开(公告)日：2023-10-19
申请号：US18211277
申请日：2023-06-18
Applicant: Special Aerospace Services, LLC
Inventor： Timothy Bulk , Christopher Hayes
Abstract: 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.
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Patent Agency Ranking
公开(公告)号：US11719194B2
公开(公告)日：2023-08-08
申请号：US17463609
申请日：2021-09-01
Applicant: ArianeGroup GmbH
Inventor： Chris Udo Maeding
Abstract: A combustion chamber for a rocket engine, the combustion chamber including a combustion chamber body enclosing a combustion chamber volume and a nozzle portion tapering in a longitudinal direction of the combustion chamber and adjoining the combustion chamber body. The combustion chamber body has at least one first portion and a second portion, wherein an inner surface of the at least one first portion facing the combustion chamber volume is closer to a cross-sectional center of the combustion chamber body than an inner surface of the second portion of the combustion chamber body. Furthermore, a additive layer manufacturing method for manufacturing such a combustion chamber is described.
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公开(公告)号：US20230220817A1
公开(公告)日：2023-07-13
申请号：US17561621
申请日：2021-12-23
Applicant: Andrew Thomas Duggleby
Inventor： Andrew Thomas Duggleby
CPC classification number: F02K9/64 , F02K9/46 , F05D2260/232 , F05D2260/213 , F05D2240/35
Abstract: A rocket engine system comprising a rocket engine, coolant and a coolant source, propellant and a propellant source, a turbopump, and a heat source. The coolant is pressurized and then heated through a heat source to a supercritical state for augmented heat transfer. The heat source may be a heat exchanger with returning coolant, or a preburner. The rocket engine system may further comprise at least one additional rocket engine with a pump to provide pressure for multiple engine. The rocket engine system may further comprise multiple turbopump shafts for independent control of propellants.
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公开(公告)号：US20190203662A1
公开(公告)日：2019-07-04
申请号：US16311820
申请日：2017-06-28
Applicant: ARIANEGROUP SAS
Inventor： Thierry PICHON , Xavier ZORRILLA , Ivan HERRAIZ , Laurent LONGUEVILLE
CPC classification number: F02K9/97 , F02K9/64 , F05D2300/10 , F05D2300/603
Abstract: A nozzle presents a longitudinal axis, includes both a combustion chamber made of metal material and presenting a downstream end, and a diverging portion made of composite material formed by a wall of conical shape extending between an upstream and a downstream end. The upstream end of the composite material diverging portion is connected to the downstream end of the combustion chamber. The nozzle further includes an annular mount made of metal material including a first portion secured to the combustion chamber and a second portion extending beyond the downstream end of the combustion chamber along the longitudinal axis. The upstream end of the composite material diverging portion is fastened to the second portion of the annular mount by a plurality of fastener members, each including a fastener bolt, each fastener bolt passing through the conically-shaped wall of the composite material diverging portion near the upstream end of the wall.
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公开(公告)号：US20180010552A1
公开(公告)日：2018-01-11
申请号：US15641995
申请日：2017-07-05
Applicant: AIRBUS DS GMBH
Inventor： Ludwig BRUMMER
CPC classification number: F02K9/64 , B23P15/008 , B23P15/26 , B23P2700/13 , F05D2230/20 , F05D2260/205 , F05D2260/213 , F23R3/005 , F23R2900/00018
Abstract: A combustion chamber suitable in particular for use in a rocket engine comprises a combustion space, a first wall enclosing the combustion space and cooling duct fins, which extend from a surface of the first wall and separate adjacent cooling ducts from one another. At least one of the cooling duct fins has at its end facing away from the surface of the first wall a bent section, which at least partially covers a cooling duct adjacent to the cooling duct fin.
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公开(公告)号：US20170335797A1
公开(公告)日：2017-11-23
申请号：US15598473
申请日：2017-05-18
Applicant: Airbus DS GmbH
Inventor： Ulrich Gotzig , Malte Wurdak , Joel Deck , Manuel Frey
CPC classification number: F02K9/56 , F02K9/425 , F02K9/58 , F02K9/64 , F02K9/95 , F05D2220/80 , F05D2240/35 , F05D2260/202 , F05D2260/2212
Abstract: A method for operating a rocket propulsion system comprises the steps of supplying oxygen to a combustion chamber, supplying hydrogen to the combustion chamber and combusting the oxygen-hydrogen mixture in the combustion chamber. The rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.
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