WO2016039993A1 - Moteur-fusée à propergol liquide ayant une chambre de post-combustion - Google Patents

Moteur-fusée à propergol liquide ayant une chambre de post-combustion Download PDF

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Publication number
WO2016039993A1
WO2016039993A1 PCT/US2015/047073 US2015047073W WO2016039993A1 WO 2016039993 A1 WO2016039993 A1 WO 2016039993A1 US 2015047073 W US2015047073 W US 2015047073W WO 2016039993 A1 WO2016039993 A1 WO 2016039993A1
Authority
WO
WIPO (PCT)
Prior art keywords
turbine
fuel
oxidizer
combustion chamber
liquid propellant
Prior art date
Application number
PCT/US2015/047073
Other languages
English (en)
Inventor
Ross HEWITT
Melvin J. Bulman
Original Assignee
Aerojet Rocketdyne, Inc.
Priority date (The priority date 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 date listed.)
Filing date
Publication date
Application filed by Aerojet Rocketdyne, Inc. filed Critical Aerojet Rocketdyne, Inc.
Publication of WO2016039993A1 publication Critical patent/WO2016039993A1/fr

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/42—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof using liquid or gaseous propellants
    • F02K9/44—Feeding propellants
    • F02K9/46—Feeding propellants using pumps
    • F02K9/48—Feeding propellants using pumps driven by a gas turbine fed by propellant combustion gases or fed by vaporized propellants or other gases
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02K—JET-PROPULSION PLANTS
    • F02K9/00—Rocket-engine plants, i.e. plants carrying both fuel and oxidant therefor; Control thereof
    • F02K9/97—Rocket nozzles

Definitions

  • liquid propellant rocket engines Numerous types of liquid propellant rocket engines are known.
  • One type of engine utilizes an expander cycle in which pressurized fuel is expanded through a turbine prior to injection into a rocket combustion chamber. The turbine drives a fuel and/or oxidizer pump.
  • Another type of engine utilizes a gas generator cycle in which fuel and oxidizer are burned in a pre-burner to generate an exhaust gas that is expanded through a turbine. The exhaust from the turbine is then dumped overboard.
  • a third type of engine utilizes a staged combustion cycle that is similar to the gas generator cycle but the exhaust from the turbine is injected into the rocket combustion chamber rather than being dumped overboard.
  • a liquid propellant rocket engine includes a combustion chamber that has a throat and a nozzle that diverges from the throat.
  • the nozzle has an afterburner combustor section that includes a fuel injector orifice.
  • the turbine inlet is fluidly coupled with an exhaust duct of a gas generator and the turbine outlet is fluidly coupled with the fuel injector orifice.
  • a further embodiment of any of the foregoing embodiments includes an oxidizer source fluidly coupled with an oxidizer injector orifice of the afterburner combustor section.
  • the gas generator is configured to produce a fuel-rich exhaust.
  • a liquid propellant rocket engine includes first and second combustion chambers having, respectively, first and second throats, and first and second nozzles that diverge, respectively, from the first and second throats.
  • a fuel source is fluidly coupled with the first combustion chamber and a gas generator that has an exhaust duct.
  • a turbine includes a turbine inlet and a turbine outlet. The turbine inlet is fluidly coupled with an exhaust duct of the gas generator and the turbine outlet is fluidly coupled with the second combustion chamber.
  • the first and second nozzles are of different sizes.
  • the turbine has a pressure ratio across the turbine inlet and the turbine outlet of approximately 2:1.
  • a method for a liquid propellant rocket engine includes delivering fuel and oxidizer to a first combustion chamber that includes a first throat.
  • the fuel and oxidizer react to generate a gas stream in a first nozzle that diverges from the first throat.
  • Fuel and oxidizer are also delivered to a gas generator that includes an exhaust duct.
  • the fuel and oxidizer react to generate a fuel-rich gas stream in the exhaust duct that expands across a turbine.
  • the expanded fuel-rich gas stream from the turbine is delivered to one of a fuel injector orifice in an afterburner combustor section of the first nozzle, or a second, different combustion chamber.
  • the fuel-rich gas stream is expanded to a pressure that is substantially equal to or greater than 50% of the pressure of the gas stream in the first combustion chamber.
  • Figure 1 illustrates an example liquid propellant rocket engine with a gas generator and afterburner combustor section downstream of the main combustion chamber.
  • Figure 2 illustrates another example liquid propellant rocket engine that is similar to the engine in Figure 1 but additionally includes an additional turbine and oxidizer pump.
  • Figure 3 illustrates another liquid propellant rocket engine that includes a gas generator and an oxidizer source feeding the afterburner combustor.
  • FIG. 1 schematically illustrates selected portions of an example liquid propellant rocket engine 20.
  • the engine 20 is mounted on a vehicle V (represented schematically) and, as will be described in further detail herein, is configured as an afterburning, gas generator cycle engine.
  • V represented schematically
  • afterburning and variations thereof refer to a burning of propellants to augment thrust generated by a main or primary combustion cycle.
  • the engine 20 includes a combustion chamber 22 that has a throat 24, which is a relatively narrow or narrowest portion of the combustion chamber 22.
  • a nozzle 26 diverges from the throat 24 and includes an afterburner combustor section 28.
  • the nozzle 26 includes a proximal section 26a and a distal section 26b with respect to proximity to the throat 24.
  • the proximal and distal section 26a/26b are taken with regard the axial length of the nozzle 26 along central axis A.
  • a line L in Figure 1 denotes the boundary between the proximal and distal sections 26a/26b.
  • the proximal section 26a is approximately 10% of the axial length of the nozzle 26.
  • the nozzle 26 includes the afterburner combustor section 28.
  • the afterburner combustor section 28 is in the proximal section 26a of the nozzle 26.
  • the afterburner combustor section 28 can be nearer (axially) to the throat 24 than to line L.
  • the afterburner combustor section 28 may be in the distal section 26b, although there may be greater performance gains if located in the proximal section 26a.
  • the afterburner combustor section 28 is generally an annular or frustoconical portion of the nozzle 26.
  • the afterburner combustor section 28 may include the portion of the wall of the nozzle 26 that includes or encompasses one or more fuel injector orifices 30 and one or more oxidizer injector orifices 32.
  • the relative locations of the fuel injector orifices 30 with respect to the locations of the oxidizer injector orifices 32 can be varied, in addition to variation in the type of injection provided by the orifices 30/32 with regard to injecting streams, fans, or sprays. In general though, the orifices 30/32 will be distributed circumferentially.
  • the fuel injector orifices 30 may be either forward or aft of the oxidizer injector orifices 32. Additionally, the orifices 30/32 may be configured to generate impinging injection flows. The orifices 30/32 additionally or alternatively can be configured to inject in a direction that is perpendicular to the central axis A, at an angle conformal to the inside surface of the nozzle 26, or at any angle in between. A further example of one or more of the orifices 30/32 can include an annular orifice.
  • the engine 20 also includes a gas generator 34.
  • the gas generator 34 includes an exhaust duct 34a that is fluidly coupled with at least one turbine 36.
  • the turbine 36 includes a turbine inlet 36a and a turbine outlet 36b.
  • the turbine inlet 36a is fluidly coupled with the exhaust duct 34a and the turbine outlet 36b is fluidly coupled with the fuel injector orifices 30.
  • the engine 20 includes at least one oxidizer source 38 and at least one fuel source 40.
  • a typical fuel can include, but is not limited to, kerosene, methane, or hydrogen.
  • a typical oxidizer can include, but is not limited to, gaseous or liquid oxygen, nitrogen tetroxide, nitrous oxide, and hydrogen peroxide.
  • the oxidizer source 38 may include one or more oxidizer pumps 38a, and the fuel source 40 may include one or more fuel pumps 40a, which are both mounted to be driven by the turbine 36.
  • the oxidizer pump 38a is fluidly coupled with the combustion chamber 22, through primary oxidizer injectors 23a, as well as the secondary (afterburner) oxidizer injector orifices 32 and the gas generator 34.
  • the fuel pump 40a is fluidly coupled with the gas generator 34 and the nozzle 26. Prior to injection into the combustion chamber 22, the fuel from the fuel pump 40a may be conveyed through internal passages in the walls of the nozzle 26, throat 24, and combustion chamber 22 for cooling, and to the primary fuel injector 22a.
  • the oxidizer source 38 delivers oxidizer and the fuel pump 40a delivers fuel.
  • a portion of the fuel is delivered to the internal passages in the nozzle 26 as a coolant, and another portion of the fuel is delivered to the gas generator 34.
  • the oxidizer is divided among the gas generator 34, the primary oxidizer injectors 23a for injection into the primary combustion chamber 22 and secondary injector orifices 32.
  • the fuel and oxidizer burn in the gas generator 34 to generate a fuel-rich exhaust that is expanded through the one or more turbines 36.
  • the expanded fuel-rich exhaust is then delivered from the turbine outlet 36b to the fuel injector orifices 30 in the afterburner combustor section 28.
  • Fuel and oxidizer are also injected through injectors 22a into the combustion chamber 22 to generate thrust through the throat 24 and nozzle 26.
  • the fuel-rich gas stream from the turbine 36 is injected through the fuel injector orifices 30 and oxidant is injected through the oxidant orifices 32 into the afterburner combustor section 28.
  • the fuel-rich gas stream and oxidant burn downstream of the throat 24 to generate additional thrust in the nozzle 26.
  • the engine 20 may be controlled with respect to certain pressures in the system.
  • the combustion chamber 22 is configured to produce a first pressure therein and the gas generator 34 and turbine 36 are configured to produce a second pressure at the turbine outlet 36b that is close to the first pressure.
  • the second pressure By controlling the second pressure to be as closer to the first pressure than typical of a conventional Gas Generator Cycle Engine, greater augmentation thrust can be produced.
  • the size of the turbine 36 can be selected with respect to a pressure ratio across the turbine inlet 36a and a turbine outlet 36b. For instance, if such a ratio is too high, such as 10: 1, the pressure at the turbine outlet 36b would be about 10% of first pressure in the combustion chamber 22, reducing the required gas generator flow and thrust efficiency. However, the use of a pressure ratio of approximately 2: 1 promotes greater thrust efficiency of the afterburner gases.
  • Figure 2 illustrates another example liquid propellant rocket engine 120 that is similar to the engine 20 but includes an oxidizer source 138 that has a first oxidizer pump 138a and a second oxidizer pump 138b.
  • the first oxidizer pump 138a is fluidly coupled with the injectors 22a of the combustion chamber 22 and also the gas generator 34.
  • the engine 120 also includes a plurality of turbines 136, which in this example includes a first turbine 136-1 and second turbine 136-2.
  • the first turbine 136-1 includes a turbine inlet 136a-l and a turbine outlet 136b-l
  • the second turbine 136-2 includes a turbine inlet 136a-2 and a turbine outlet 136b-2.
  • the first turbine 136-1 is coupled to, and drives, the fuel pump 40a and the first oxidizer pump 138a.
  • the second turbine 136-2 is coupled to, and drives, the second oxidizer pump 138b.
  • the fuel-rich exhaust gas stream is expanded through the first turbine 136-1, but in this example is then also expanded through the second turbine 136-2 to drive the second oxidizer pump 138.
  • the second oxidizer pump 138 serves to deliver the oxidizer to the oxidizer injector orifices 32.
  • FIG. 3 illustrates another example liquid propellant rocket engine 220.
  • a second, different combustion chamber 50 serves for the afterburning.
  • the second combustion chamber 50 includes a second throat 52 and a second nozzle 54 that diverges from the second throat 52.
  • the first and second combustions chambers 22/50, first and second throats 24/52, and first and second nozzles 26/54 are separate and distinct from each other and thus discharge separate or uncommon effluent streams.
  • the first and second nozzles 26/54 are of different sizes, with the nozzle 54 being generally smaller.
  • the first and second combustions chambers 22/50 and the first and second throats 24/52 may also be of different sizes than each other.
  • the oxidizer pump 38 delivers oxidizer to the gas generator 34, the combustion chamber 22, and the combustion chamber 50.
  • the fuel-rich gas stream from the one or more turbines 36 is delivered, in addition to oxidizer, into the second combustion chamber 50, which serves as an afterburner combustor, to provide afterburning thrust to augment the thrust generated by the main or primary first combustion chamber 22.
  • the gas generator 34 and the one or more turbines 36 can be configured to provide a pressure at the turbine outlet 36b that is equal to or greater than 50% of the pressure in the combustion chamber 22.
  • the above examples and figures also embody exemplary methods for a liquid propellant rocket engine.
  • the method includes delivering fuel and oxidizer to the (first) combustion chamber 22, delivering fuel and oxidizer to the gas generator 34, expanding the fuel-rich gas stream across the turbine 36/136-1/136-2 and, depending on which of the engines 20/120/220, delivering the expanded fuel-rich gas stream from the turbine to one of: (i) the fuel injector orifices 30, or (ii) the second combustion chamber 50.
  • the method includes expanding the fuel-rich gas stream to a pressure that is substantially equal to approximately 50% of the pressure of the gas stream in the (first) combustion chamber 22.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Abstract

L'invention concerne un moteur-fusée à propergol liquide qui comprend une chambre de combustion qui comporte un col. Une buse se sépare du col. La buse comprend une section chambre de post-combustion qui comporte un orifice d'injecteur de carburant. Un générateur de gaz comprend un conduit d'échappement et au moins une turbine, qui comporte une entrée de turbine et une sortie de turbine, est présente. L'entrée de turbine est en communication fluidique avec le conduit d'échappement, et une sortie de turbine est en communication fluidique avec l'orifice d'injecteur de carburant.
PCT/US2015/047073 2014-09-12 2015-08-27 Moteur-fusée à propergol liquide ayant une chambre de post-combustion WO2016039993A1 (fr)

Applications Claiming Priority (6)

Application Number Priority Date Filing Date Title
US201462049855P 2014-09-12 2014-09-12
US62/049,855 2014-09-12
US201462062094P 2014-10-09 2014-10-09
US62/062,094 2014-10-09
US201462085538P 2014-11-29 2014-11-29
US62/085,538 2014-11-29

Publications (1)

Publication Number Publication Date
WO2016039993A1 true WO2016039993A1 (fr) 2016-03-17

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Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112555056A (zh) * 2020-12-02 2021-03-26 西安航天动力研究所 补燃循环液体发动机核心系统热试验装置及参数协调方法
CN114991994A (zh) * 2022-05-24 2022-09-02 南京航空航天大学 一种固体火箭双冲压组合发动机及工作方法
RU2789943C1 (ru) * 2022-06-21 2023-02-14 Владимир Федорович Петрищев Жидкостный ракетный двигатель с форсажем
CN117329025A (zh) * 2023-12-01 2024-01-02 陕西天回航天技术有限公司 一种涡轮排气冲压增推组合循环发动机及航天飞行器

Citations (5)

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Publication number Priority date Publication date Assignee Title
GB1210601A (en) * 1967-02-11 1970-10-28 Mtu Muenchen Gmbh Rocket propulsion system
RU2065068C1 (ru) * 1994-08-03 1996-08-10 Конструкторское бюро химавтоматики Экспериментальный жидкостный ракетный двигатель с дожиганием
US6227486B1 (en) * 1999-05-28 2001-05-08 Mse Technology Applications, Inc. Propulsion system for earth to orbit vehicle
RU2204046C2 (ru) * 2000-02-15 2003-05-10 Открытое акционерное общество "Самарский научно-технический комплекс им. Н.Д. Кузнецова" Жидкостный ракетный двигатель с дожиганием
US8250853B1 (en) * 2011-02-16 2012-08-28 Florida Turbine Technologies, Inc. Hybrid expander cycle rocket engine

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1210601A (en) * 1967-02-11 1970-10-28 Mtu Muenchen Gmbh Rocket propulsion system
RU2065068C1 (ru) * 1994-08-03 1996-08-10 Конструкторское бюро химавтоматики Экспериментальный жидкостный ракетный двигатель с дожиганием
US6227486B1 (en) * 1999-05-28 2001-05-08 Mse Technology Applications, Inc. Propulsion system for earth to orbit vehicle
RU2204046C2 (ru) * 2000-02-15 2003-05-10 Открытое акционерное общество "Самарский научно-технический комплекс им. Н.Д. Кузнецова" Жидкостный ракетный двигатель с дожиганием
US8250853B1 (en) * 2011-02-16 2012-08-28 Florida Turbine Technologies, Inc. Hybrid expander cycle rocket engine

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN112555056A (zh) * 2020-12-02 2021-03-26 西安航天动力研究所 补燃循环液体发动机核心系统热试验装置及参数协调方法
CN114991994A (zh) * 2022-05-24 2022-09-02 南京航空航天大学 一种固体火箭双冲压组合发动机及工作方法
CN114991994B (zh) * 2022-05-24 2023-03-28 南京航空航天大学 一种固体火箭双冲压组合发动机及工作方法
RU2789943C1 (ru) * 2022-06-21 2023-02-14 Владимир Федорович Петрищев Жидкостный ракетный двигатель с форсажем
RU2813564C1 (ru) * 2023-07-31 2024-02-13 Владимир Федорович Петрищев Способ работы жидкостного ракетного двигателя с форсажем
CN117329025A (zh) * 2023-12-01 2024-01-02 陕西天回航天技术有限公司 一种涡轮排气冲压增推组合循环发动机及航天飞行器
CN117329025B (zh) * 2023-12-01 2024-02-23 陕西天回航天技术有限公司 一种涡轮排气冲压增推组合循环发动机及航天飞行器

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