• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Rocket engine (1) having a main combustion chamber (5) and a main nozzle; two propellant supply pumps (8a, 13) arranged on engine propellant supply circuits for supplying the main combustion chamber; two pressurizing pumps (10,11) for pressurizing the propellants; and an electric generator (20). The electric generator (20) comprises a gas generator (22), an electricity generating turbine (24) configured to be driven by gases exiting the gas generator, and an alternator (26) mechanically coupled to the shaft of the electricity generation turbine. The rocket engine is configured to allow the propellant gas generator to be fed by the two pressurizing pumps (10, 11). The power supply of the electricity generator (24) is independent of the rotation of the feed pumps (8a, 13).
    公开号:FR3062171A1
    申请号:FR1700054
    申请日:2017-01-23
    公开日:2018-07-27
    发明作者:Didier Vuillamy；Emmanuel Edeline；Francois Lassoudiere
    申请人:Airbus Safran Launchers SAS；
    IPC主号:
    专利说明:

    © Holder (s): AIRBUS SAFRAN LAUNCHERS SAS Simplified joint-stock company.
    O Extension request (s):
    © Agent (s): CABINET BEAU DE LOMENIE.
    FR 3 062 171 - A1 (54) MOTOR-ROCKET.
    (57) Rocket engine (1) comprising a main combustion chamber (5) and a main nozzle; two propellant supply pumps (8a, 13) arranged on engine propellant supply circuits for supplying the main combustion chamber;
    two pressurization pumps (10,11) for pressurizing the propellants; and an electric generator (20).
    The electric generator (20) includes a gas generator (22), an electricity generating turbine (24) configured to be driven by gases exiting the gas generator, and an alternator (26) mechanically coupled to the shaft of the electricity generation turbine.
    The rocket engine is configured so as to allow the gas generator to be supplied with propellants by the two pressurization pumps (10,11).
    The supply of the electricity generator (24) is independent of the rotation of the supply pumps (8a, 13).

    FIELD OF THE INVENTION
    The invention relates to a rocket engine and, more specifically, to the production of electrical energy in a rocket engine.
    It relates in particular, but not exclusively, to rocket engines comprising a propellant supply circuit comprising a turbopump which is used to put the flow of propellant circulating in the circuit under pressure before the injection of the latter into the combustion chamber of the motor.
    More particularly, the invention relates to rocket motors of the type indicated above whose thermodynamic cycle is of the "expander" type.
    A propellant supply circuit for an “expander” type motor connects the outlet of the turbopump pump with the inlet of the turbopump turbine through a heat exchanger configured to heat the propellant with heat generated in the combustion chamber to activate the turbine of the turbopump by expansion of the first propellant after its heating. The gases leaving the turbine of the turbopump are then directed to the combustion chamber.
    BACKGROUND OF THE INVENTION
    We know, for example from document FR2991391, different sources of electrical energy usable in rocket engines of the type indicated above.
    In such engines, an electric generator can for example be arranged on the shaft of the turbine of the turbopump to produce electricity.
    However, this arrangement has the disadvantage that the electric generator produces electricity only when the turbopump is driven in rotation; it therefore remains unusable for the start-up phases.
    Alternatively, it is also possible to integrate a fuel cell or batteries into the rocket engine. These solutions induce increased complexity and / or mass, and therefore remain quite unsatisfactory.
    OBJECT AND SUMMARY OF THE INVENTION
    The objective of the invention is therefore to remedy the drawbacks indicated above, and to propose a rocket engine comprising a flexible source of electricity, which can be used to supply the rocket engine with electricity preferably during the whole flight phases.
    This objective is achieved thanks to a rocket engine comprising a main combustion chamber;
    a main nozzle for ejecting the combustion gases produced in the main combustion chamber;
    two propellant feed pumps disposed on engine propellant feed circuits and configured to pressurize the propellants to provide power to the main combustion chamber;
    two pressurization pumps to pressurize the propellant tanks used to power the engine;
    an electric generator; the electric generator comprising:
    - a gas generator configured to be supplied with a mixture of propellants to cause combustion;
    - an electricity generation turbine configured to be driven by combustion gases from the gas generator; and
    - an alternator mechanically coupled to a shaft of the electricity generation turbine;
    the rocket engine being configured so as to allow the gas generator to be supplied with propellants by the two pressurization pumps, and so as to allow the electricity generator to be supplied independently of the rotation of the supply pumps.
    In a manner known per se, the engine is generally connected to two fuel tanks. The pressurization pumps are then used to pressurize the propellant tanks used to supply the engine, that is to say to maintain these tanks at an adequate pressure in particular during the operating phases of the rocket engine.
    Each of the pressurization pumps can be placed (or not) inside the tank which it serves to maintain under pressure. Placing a pressurization pump in the tank allows, advantageously for cryogenic propellants, to dispense with a cooling phase prior to starting the rocket engine.
    The electric generator supplies electricity to the rocket engine.
    The electricity it produces can notably be used to power one or more of the supply pumps (if these are electric). These pumps can then pump the propellant (s) so as to supply the combustion chamber. It can also be used to power pressurization pumps, or more generally any electrical component of the engine, for example solenoid valves, etc.
    As the rocket motor is configured so that the power supply to the electricity generator is independent of the rotation of the supply pumps, the electricity generator can operate independently of the fact that the supply pump (s) are running or not ; it can thus operate when they rotate, or when they are stopped. The electric generator can therefore operate independently of the supply of propellants to the engine, and therefore independently of the engine operating regime. Consequently, the rocket engine benefits from a supply of electricity which can be used at any time, whatever the flight phase.
    That said, it is of course necessary to adopt adequate provisions to allow the supply of the gas generator during the brief start-up period of the electric generator: in fact, before it has started, it cannot drive the pressurization pumps . Two solutions in particular can be envisaged: one can either provide a small battery, sized to briefly actuate the pressurization pumps until the electric generator starts; or else, the gas pressure in the propellant tanks may be sufficient to allow the electric generator to start.
    Each of the propellant supply pumps can be an electric pump or a turbopump. For example, one of the supply pumps may be an electric pump, and the electric generator is configured to supply the latter with electricity.
    One of the supply pumps can also be a pump of a turbopump arranged downstream of one of the pressurization pumps on a first supply circuit allowing the supply of propellant to the main combustion chamber of the rocket engine.
    In a variant of this latter embodiment, the rocket motor can be of the so-called “expander” type, that is to say that the first supply circuit connects the output of the pump of the turbopump with the input of a turbine of the turbopump through a heat exchanger configured to heat the first propellant with heat generated in a combustion chamber of the rocket engine to actuate the turbine of the turbopump by expansion of the first propellant after its heating. In this embodiment, the gases escaping from the turbine of the turbopump are completely reinjected into the main combustion chamber of the rocket engine.
    In general, the gases escaping from the electricity generation turbine can be discharged in different ways to the outside. In particular, these gases can be discharged into the main combustion chamber, and / or into an auxiliary nozzle used to eject these gases outside. In the latter case, preferably the rocket engine comprises an auxiliary nozzle, to which is connected an exhaust orifice of the electricity generation turbine. The expression “auxiliary nozzle” designates here a nozzle distinct from the main nozzle of the rocket engine and of small size compared to the latter. Advantageously, the pressure of the gases leaving the electricity generation turbine can then be very low. Thanks to the ejection of gases into the auxiliary nozzle, advantageously the turbine can have a high expansion ratio, and thus transmit high mechanical power to the alternator.
    On the other hand, when the rocket engine comprises an auxiliary nozzle to which is connected an exhaust orifice of the electricity generating turbine, and a supply circuit of the rocket engine comprises a turbopump, advantageously the auxiliary nozzle can also recover energy from the gases released by the turbine of the turbopump.
    In this case, provision may be made for an exhaust orifice from the turbine of the turbopump to be connected to a duct which has at least two branches, namely a first branch serving to conduct the exhaust gases from the turbine. from the turbopump to the combustion chamber, and a second branch connected to an inlet orifice of the electricity generation turbine. Advantageously, the electric generator then makes it possible to produce electricity from the gases discharged by the turbine of the turbopump. In addition, this may possibly make it possible to stop the gas generator: electricity is then produced by turning the electricity generation turbine only from the gases discharged by the turbine of the turbopump.
    The first supply circuit can be arranged in different ways.
    In an embodiment indicated above, the rocket engine is of the so-called “expander” type. In this case, the rocket engine can be configured so as to lead all of the gases escaping from the turbine of the turbopump into the combustion chamber. For example, the exhaust port of the turbopump turbine can be connected to a turbopump turbine exhaust duct, which is used to conduct gases escaping from the turbopump turbine only to the chamber combustion. The first circuit is then of the closed cycle type, the flow of first propellant heated with the heat generated in the combustion chamber being fully reinjected therein.
    This embodiment does not, however, exclude that the electricity generation turbine further comprises an exhaust duct which opens into the exhaust duct of the turbopump turbine, and thus makes it possible to reject the exhaust gases therein. by the electricity generation turbine.
    Furthermore, as previously indicated, the rocket engine may include an auxiliary nozzle.
    In one embodiment, the auxiliary nozzle is used only to reject the combustion gases produced by the gas generator. However, alternatively (or in addition) the auxiliary nozzle can be used to reject other gases. Thus, when a feed pump is a turbopump, the rocket motor can thus in particular comprise a conduit for conducting the gases escaping from the turbine of the turbopump (for an expander circuit, it is a propellant in vapor phase) to the auxiliary nozzle.
    In this case, the first circuit can then be of the so-called “full bleed” or “partial bleed” type, depending on whether it is configured to conduct all or part of the gases escaping from the turbine of the turbine into the auxiliary nozzle. turbopump.
    In the case where the first circuit is of the “partial bleed” type, the first circuit is configured to conduct part of the gases escaping from the turbine in said auxiliary nozzle, and part in the combustion chamber. The rocket engine control unit is normally then intended to control, by means of controllable valves, the relative proportions of the gases escaping from the turbine which are directed respectively to the auxiliary nozzle and to the combustion chamber, depending of the flight phase and the values of different parameters.
    Consequently, the following different arrangements are possible in the case where the rocket engine comprises an auxiliary nozzle:
    - an exhaust port of the turbine of the turbopump is connected to the auxiliary nozzle;
    - the exhaust port of the turbopump turbine is connected to a duct which has at least two branches, namely a first branch used to conduct the exhaust gases from the turbopump turbine to the combustion, and a second branch connected to the auxiliary nozzle and / or to an outlet or inlet of the electricity generation turbine.
    The second branch of the conduit connected to the outlet orifice of the electricity generation turbine allows, if necessary, that the outlet gases from this turbine are evacuated via the main combustion chamber of the rocket engine.
    All or part of the improvements indicated above in relation to the first circuit supplying the rocket engine with the first propellant can of course also possibly be implemented for any other supply circuit used for supplying propellant (s) of the rocket engine.
    On the other hand, as has been explained previously, the gas generator can possibly be used only to produce electricity for the rocket engine.
    However, the present invention does not exclude that the energy of the combustion gases produced by the gas generator is used for other uses. For example, this energy can be used to pressurize one or more propellant tanks.
    Thus in one embodiment, the rocket engine further comprises a heat exchanger capable of heating and / or vaporizing at least in part one of the propellants by drawing heat from the combustion gases produced by the gas generator.
    The heat exchanger and the associated circuit then constitute a pressurization device capable of ensuring the maintenance of a predetermined pressure in the first tank.
    Although in this embodiment, it is only indicated that the first tank is provided with such a pressurization system, it is however understood that more generally, within the framework of the invention, any propellant tank may or may not (and therefore possibly, several tanks) can be equipped with such a pressurization system, in order to exploit the energy produced by the combustion of gases from the gas generator.
    Preferably, the invention is implemented in a rocket engine configured to generate a maximum thrust of less than 100 kN. It can however be implemented in engines developing a higher thrust.
    BRIEF DESCRIPTION OF THE DRAWINGS
    The invention will be clearly understood and its advantages will appear better on reading the detailed description which follows, of several embodiments shown by way of nonlimiting examples. The description refers to the accompanying drawings in which:
    - Figure 1 is a schematic view of a rocket engine according to a first embodiment of the invention;
    - Figure 2 is a schematic view of a rocket engine according to a second embodiment of the invention;
    - Figure 3 is a schematic view of a rocket engine according to a third embodiment of the invention; and
    - Figure 4 is a schematic view of a rocket engine according to a fourth embodiment of the invention.
    DETAILED DESCRIPTION OF THE INVENTION
    The invention will now be illustrated by four embodiments showing rocket engines having different types of combustion cycle.
    The rocket engines presented are liquid hydrogen and oxygen engines, but the invention can be implemented with other propellants. For example, the fuel can also be methane, ethanol or kerosene.
    FIG. 1 illustrates a rocket engine 1 comprising a combustion chamber 5 and an electric generator 20 for supplying electricity to the rocket engine 1, according to a first embodiment.
    The rocket engine 1 comprises a tank 2 containing hydrogen in the liquid state, a tank 3 containing oxygen in the liquid state, a hydrogen circuit 4 connected to the tank 2 to supply the combustion chamber 5 in hydrogen, and an oxygen circuit 6 connected to the tank 3 to supply the combustion chamber 5 with oxygen.
    In this embodiment, the hydrogen circuit 4 comprises a pressurization pump 10, an inlet valve 7, a turbopump 8 with a feed pump 8a and a turbine 8b mechanically coupled, and a heat exchanger 9 formed in the walls of the combustion chamber 5 so as to transfer heat from the combustion chamber 5 to the hydrogen during its circulation through the heat exchanger 9. The heat exchanger 9 is located, in the first circuit 4, downstream of the pump 8a and upstream of the turbine 8b. Thus, the transfer of heat in the exchanger 9 simultaneously contributes to cooling the walls of the combustion chamber 5 and to vaporizing the liquid hydrogen between the pump 8a and the turbine 8b. The expansion of the hydrogen in the turbine 8b actuates the turbopump 8. Thus, this hydrogen circuit 4 of the first embodiment operates according to an "expander" cycle.
    This hydrogen circuit 4 also includes a passage 15 for bypassing or bypassing the turbine 8b, with a bypass valve 16. By this passage 15, the flow of hydrogen leaving the heat exchanger 9 can be injected into the combustion chamber 5 without passing through the turbine 8b.
    The pressurization pump 10 is an electric pump immersed in liquid hydrogen in the first tank 2 and is used to pump the hydrogen through the circuit 4, thus ensuring the feeding of the pump 8a and preventing cavitation phenomena. This pump 10 and the inlet valve 7 can be integrated in the same module inside the tank 2 of liquid hydrogen, so as to simplify the assembly and limit its size.
    The oxygen circuit 6 mainly comprises a pressurization pump 11, an inlet valve 12 and a supply pump 13. The pressurization pump 11 is an electric pump immersed in the tank 3 of liquid oxygen which is used to pump liquid oxygen via the circuit 6 to the suction orifice of the second electric pump 13. The supply pump 13 is a high-power pump, in particular of much higher power than that of the pressurization pump 11 In a variant of the embodiment presented, the inlet valve 12 can optionally be integrated in the same module as the pressurization pump 11 and thus placed inside the tank 3 of liquid oxygen. The feed pump 13 can be placed outside or inside the tank 3.
    Unlike the hydrogen circuit 4, the liquid oxygen circuit 6 does not include a turbopump; thus, the rocket motor 1 advantageously comprises only one turbopump (8), or at least, a single turbopump of high power (several hundred kilowatts).
    For the supply of electricity to the supply pump 13 and (except during the initial start-up phase) of the pressurization pumps 10 and 11, the supply device comprises an electric generator 20.
    This electric generator 20 comprises a gas generator 22, an electricity generation turbine 24, and an alternator 26 mechanically coupled with a shaft of the turbine 24.
    The gas generator 22 is configured to be supplied with a mixture of propellants independently of the state of rotation of the pumps 8a and 13. For this purpose, it is connected to branches 44 and 46 respectively of circuits 4 and 6 for its supply of hydrogen and oxygen. Each of these branches includes a controllable valve (valves 30, 32); these valves 30 and 32 thus make it possible to control the operation of the gas generator 22 and consequently, of the electric generator 20. These 5 branches 44 and 46 are connected to the circuits 4 and 6 respectively upstream of the pump 8a and of the pump 13 Advantageously, the gas generator 22 can be supplied with propellants via the branches 44 and 46, whether the pumps 8a and 13 are in operation or stopped. The operation of the gas generator 22 and therefore of the electric generator 20 only requires operating the low-power electric pumps 10 and 11, but does not require operating the turbopump 8a or the pump 13, both of significantly higher power high.
    In the gas generator 22, the hydrogen is burned by oxygen. The combustion gases produced escape through an exhaust circuit which passes through the electricity generation turbine 24. Under the effect of the pressure of these gases, the electricity generation turbine 24 is driven in rotation.
    The energy communicated to the electricity generation turbine 24 by the combustion gases is then used to produce electricity thanks to the alternator 26, which is coupled to the shaft of the electricity generation turbine 20 24. The energy produced by the combustion gases therefore allows the electric generator 20 to produce electricity.
    Downstream of the turbine 24, the gases escaping from the turbine 24 are reinjected into the hydrogen circuit downstream of the turbine 8b of the turbopump 8 (at a point A, Fig.l); the first supply circuit 4 is therefore arranged in a closed cycle.
    A controllable valve 50 is provided on the gas evacuation circuit of the turbine 24. In the case where the engine is running at full power, this valve makes it possible to prevent the gases leaving the turbine 8b from going back into the turbine 24 (in the opposite direction to the desired direction of traffic).
    The turbine 24 operates independently of the turbopump 8 and of the electric pump 13. In particular, it does not serve to drive the pump 8a of the turbopump 8 in rotation.
    The electric generator 20 and the turbopump 8 can therefore be dimensioned independently of one another.
    Advantageously, the electric generator 20 can be implemented independently of the operating regime of the rocket engine 1. It can
    therefore be implemented during all phases of flight, as well as before and after the flight.
    To ensure the pressurization of the oxygen tank 3, the rocket engine further includes a heat exchanger 40. This is arranged on the duct 45 leading the (hot) combustion gases produced by the gas generator 22 to the turbine 24.
    A bypass duct 42 is arranged as a bypass on the bypass 46 supplying the gas generator 22 with oxygen. The bypass duct 42 makes it possible to take liquid oxygen and to heat and optionally at least partially vaporize it in the heat exchanger 40, then to reinject the oxygen thus heated and possibly vaporized at the top of the tank 3 The oxygen flow rate in the duct 42 is regulated by a controlled valve 48.
    The heat exchanger 40 in combination in particular with the duct 42 and the valve 48 constitutes a pressurization device making it possible to maintain a predetermined pressure in the oxygen tank 3. (The heat exchanger 40 could possibly, in a different embodiment, be arranged to heat hydrogen and thus also ensure the pressurization of the hydrogen tank).
    In the rocket engine 1, the pressurization pumps 10, 11, the inlet valves 7, 12, the electric generator 20 (in particular the valves 30 and 32), the bypass valve 16, the controlled valve 48, etc. , are connected to a control unit (not shown) for controlling the rocket motor
    1.
    On the other hand, for safety, a valve 53 is placed on the hydrogen supply circuit 4 of the pump 8, downstream of the branch point T to which the branch 44 of hydrogen supply to the gas generator 22 is connected. .
    This valve 53 is a controllable valve which serves to prevent any rise in gas through the pump 8 to the reservoir 2. It is closed in particular when it is detected that the pump 8 is stopped, to prevent any rise in the gases discharged by the gas generator 22 to the tank through the pump 8, in case the pressure at the outlet of the turbine 24 is higher than the pressure of the tank 2.
    Finally, the rocket engine 1 includes a battery, not shown. This ensures the operation of the pressurization pumps during the start-up phase of the electric generator 20, before the latter produces sufficient electric power to activate the pressurization pumps: during this phase, it is this battery which actuates the pressurization.
    Although, in this first embodiment, the circuit 4 operates according to an “expander” cycle, in alternative embodiments the turbopump can be actuated differently.
    Rocket motors 1 according to a second and a third embodiment are thus illustrated respectively in FIGS. 2 and 3. Most of the elements of these rocket motors 1 are identical or equivalent to those of FIG. 1 and consequently receive the same reference signs.
    The rocket engine 1 according to the second embodiment (Figure 2) differs from that of the first embodiment in that the gases escaping from the turbine 24 are not redirected to the hydrogen circuit 4, but to a nozzle auxiliary 28.
    The rocket engine 1 according to the third embodiment (Figure 3) is an intermediate embodiment between the first and second embodiments.
    In fact, in this third embodiment, the gases escaping from the turbine 24 can be:
    - reinjected into the hydrogen circuit downstream of the turbine 8b of the turbopump 8 (point A), and / or
    - directed towards an auxiliary nozzle 28.
    The rocket engine control unit 1 directs the gases escaping from the turbine 24 either into the hydrogen circuit 4 or into the auxiliary nozzle 28, depending on the flight phase, of the parameters of the engine parameters. , etc., by means of controlled valves 50, 52, arranged on the pipes serving respectively to reinject these gases into the hydrogen circuit downstream of the turbine 8b (at point A), and / or to direct them towards the auxiliary nozzle 28. The exhaust duct of the electricity generation turbine 24 is therefore divided into two branches, namely a first branch opening into the duct serving to conduct the exhaust gases from the turbine 8b to the combustion chamber, and a second branch opening into the auxiliary nozzle 28.
    The invention can also be implemented in a fourth embodiment (Fig. 4). This embodiment is almost identical to the third embodiment, except for two differences.
    The first difference is that the gas outlet pipe leaving the turbine 8b, is not connected (from point A, Fig. 3 or 4) to the gas outlet pipe leaving the electricity generation turbine 24 , but to the supply duct of the electricity generation turbine 24. The electricity generation turbine is therefore interposed on the duct used to conduct the gases escaping from the turbine 8b to the auxiliary nozzle 28. It can thus be supplied by the gases discharged by the turbine 8b and can therefore produce electricity from the pressure applied by these gases.
    The electric generator 20 operates independently of the supply pump 8a (and of course of the pump 13), insofar as whatever the operating regime of the pump 8a, it can operate by being supplied by the generator gas 22, in particular when the pipe going from the turbine 8b the turbine 24 is formed by the valve 50.
    The second difference compared to the third embodiment of the invention is that the hydrogen circuit 4 does not include the valve 53. Instead, a controllable valve 54 is placed on the outlet pipe of the turbine 8b of the turbopump 8, upstream of the connection point of the bypass or bypass passage 15 of the turbine 8b.
    This valve 54 has the same function as the valve 53 described above and is controlled in the same way, for the same purpose of preventing a return to the hydrogen tank 2 of the gases discharged by the gas generator 22.
    Although the present invention has been described with reference to specific exemplary embodiments, it is obvious that various modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. For example, the valve 53 can also be replaced by the valve 54 in the first and third embodiment.
    In addition, individual features of the various embodiments discussed can be combined in additional embodiments. Therefore, the description and the drawings should be considered in an illustrative rather than restrictive sense.
    权利要求:
    Claims (11)
    [1" id="c-fr-0001]
    1. Rocket engine (1) comprising a main combustion chamber (5);
    a main nozzle for ejecting the combustion gases produced in the main combustion chamber;
    two propellant feed pumps (8a, 13) disposed on engine propellant feed circuits and configured to pressurize the propellants to provide power to the main combustion chamber;
    two pressurization pumps (10,11) for pressurizing the propellant tanks used to power the engine; and an electric generator (20);
    the electric generator (20) comprising:
    - a gas generator (22) configured to be supplied with a mixture of propellants to cause combustion;
    - an electricity generation turbine (24) configured to be driven by combustion gases from the gas generator; and
    - an alternator (26) mechanically coupled to a shaft of the electricity generation turbine;
    the rocket engine being characterized in that the rocket engine is configured so as to allow the gas generator to be supplied with propellants by the two pressurization pumps (10,11), and is configured so as to allow a supply of the electricity generator (24) independently of the rotation of the feed pumps (8a, 13).
    [2" id="c-fr-0002]
    2. Rocket motor according to claim 1, in which one of the supply pumps is an electric pump (13), and the electric generator (20) is configured to supply the latter with electricity.
    [3" id="c-fr-0003]
    3. Rocket engine according to claim 1 or 2, comprising an auxiliary nozzle (28), to which is connected an exhaust orifice of the electricity generation turbine (24).
    [4" id="c-fr-0004]
    4. Rocket engine according to any one of claims 1 to 3, in which one of the feed pumps is a pump (8a) of a turbopump (8) disposed downstream of one of the pressurization pumps on a first supply circuit (4) allowing the supply of propellant to the main combustion chamber of the rocket engine.
    [5" id="c-fr-0005]
    5. Rocket engine according to claims 3 and 4, wherein a conduit connected to an exhaust port of the turbine (8b) of the turbopump has at least two branches, namely a first branch for conducting the gases exhaust of the turbine (8b) from the turbopump to the combustion chamber, and a second branch connected to an inlet orifice of the electricity generation turbine (24).
    [6" id="c-fr-0006]
    6. Rocket engine according to claim 4 or 5, wherein said first supply circuit (4) connects the outlet of the pump (8a) of the turbopump (8) with the inlet of a turbine (8b) of the turbopump through a heat exchanger (9) configured to heat the first propellant with heat generated in a combustion chamber (5) of the rocket engine to actuate the turbine (8b) of the turbopump by expansion of the first propellant after its heater.
    [7" id="c-fr-0007]
    7. Rocket engine according to any one of claims 4 to 6, comprising an auxiliary nozzle (28), and in which an exhaust orifice of the turbine (8b) of the turbopump is connected to the auxiliary nozzle (28) .
    [8" id="c-fr-0008]
    8. Rocket engine according to any one of claims 4 to 7, in which a duct connected to an exhaust orifice of the turbine (8b) of the turbopump comprises at least two branches, namely a first branch used for driving the exhaust gases from the turbine (8b) from the turbopump to the combustion chamber, and a second branch connected to an auxiliary nozzle (28) and / or to an outlet orifice of the turbine for generating electricity (24).
    [9" id="c-fr-0009]
    9. Rocket engine according to claim 4, configured so as to conduct all of the gases escaping from the turbine (8b) of the turbopump into the combustion chamber (5).
    [10" id="c-fr-0010]
    10. Rocket engine according to any one of claims 1 to 7, further comprising a heat exchanger (40) capable of heating and / or at least partially vaporizing one of the propellants by drawing heat from combustion produced by the gas generator (22).
    [11" id="c-fr-0011]
    11. Rocket engine according to any one of claims 1 to 10, 5 configured to generate a maximum thrust of less than 100 kN.
    类似技术:
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    同族专利:
    公开号 | 公开日
    FR3062171B1|2021-04-09|
    EP3571385B1|2021-04-28|
    EP3571385A1|2019-11-27|
    WO2018134541A1|2018-07-26|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    FR2991391A1|2012-05-30|2013-12-06|Snecma|Feeding device for feeding cryogenic propellant e.g. liquid oxygen, into propellant chamber of rocket engine, has tank containing cryogenic propellant, and electric pump connected within tank for pumping propellant through feeding circuit|
    US20140260186A1|2013-03-15|2014-09-18|Patrick R.E. Bahn|Rocket engine systems with an independently regulated cooling system|
    US20150288309A1|2014-04-02|2015-10-08|Hamilton Sundstrand Corporation|Systems utilizing a controllable voltage ac generator system|RU2729310C1|2019-08-04|2020-08-05|Андрей Владимирович Иванов|Liquid-propellant engine|
    RU2760956C1|2020-11-10|2021-12-01|Акционерное общество "КБхиммаш им. А.М. Исаева"|Liquid rocket engine with an electric pump supply system|
    CN109281774B|2018-12-03|2019-12-06|上海空间推进研究所|Electric pump pressure type liquid oxygen methane space propulsion system|
    法律状态:
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    2018-07-27| PLSC| Publication of the preliminary search report|Effective date: 20180727 |
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1700054A|FR3062171B1|2017-01-23|2017-01-23|ROCKET ENGINE|
    FR1700054|2017-01-23|FR1700054A| FR3062171B1|2017-01-23|2017-01-23|ROCKET ENGINE|
    EP18704278.3A| EP3571385B1|2017-01-23|2018-01-19|Rocket engine|
    PCT/FR2018/050142| WO2018134541A1|2017-01-23|2018-01-19|Rocket engine|
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