• 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:
    The invention relates to the field of rocket motors, and in particular to a method and circuit for regulating a rocket motor (1) comprising at least one combustion chamber (3) and a first power supply circuit (10). liquid propellant with a first pump (12) for a first liquid propellant and a first turbine (14) for actuating the first pump (12), a first supply valve (15) and a regulating device of the first turbine (14). In these method and control circuit an opening setpoint (DVCH) of the first supply valve (15) is calculated from an external setpoint (Cext), according to an open loop control law, a setpoint for the control device of the first turbine (14) is calculated from said external setpoint (Cext) and at least one feedback value, according to a closed-loop control law, and the first supply valve (15) ) and the regulating device of the first turbine (14) are controlled according to the respective instructions.
    公开号:FR3042227A1
    申请号:FR1502103
    申请日:2015-10-08
    公开日:2017-04-14
    发明作者:Manuel Klein；David Hayoun；Gonidec Serge Le；Sebastien Reichstadt
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    Background of the invention
    The present invention relates to the field of rocket motors with at least one liquid propellant and more precisely their regulation.
    The term "rocket engine" in this sense, any anaerobic jet engine, capable of generating a thrust by the relaxation and acceleration of a high enthalpy gas in a nozzle, in particular a convergent-divergent nozzle. Among the rocket engines, there are in particular chemical rocket engines, wherein the high enthalpy gas is generated by a chemical reaction in at least one combustion chamber upstream of the nozzle. These chemical rocket engines can be ergol propellant or solid propellant, the high enthalpy gas is generated by chemical reaction from at least one propellant. Thus, the term "rocket engine with at least one liquid propellant" means a chemical rocket engine with a single liquid propellant (monoergol), with several liquid propellants (multigol) or hybrid. In such a rocket engine with at least one liquid propellant it is possible in principle to regulate the high enthalpy gas generation, and therefore the thrust, through the supply of the propellant combustion chamber or liquid propellants. Several alternatives are known for providing this supply, including in particular the turbopump feed, in which the rocket engine combustion chamber is fed with at least one first propellant through a first liquid propellant feed circuit with a first pump. for supplying the combustion chamber with a first liquid propellant and a first turbine for actuating the first pump.
    In the field of rocket motors, it is more and more sought to regulate the thrust over a wide range of values, for example from 20 to 100% of their nominal thrust. In space launchers, for example, this makes it possible to increase the versatility of launchers, especially for multiple launches, and thus improve their economic efficiency. On satellites or orbital transfer vehicles, such an enlarged thrust range makes it possible to extend the lifetime of the machines, with also very favorable economic benefits.
    However, this regulation typically uses relatively complex control methods and circuits, involving the control of a large number of valves and other devices for regulating the supply of the combustion chamber to at least one liquid propellant. This may require a significant calculation effort, difficult to provide with the limited means embedded on such a machine, and also makes the regulation vulnerable to failures of each of these multiple control devices.
    Object and summary of the invention
    The present disclosure therefore aims to remedy these drawbacks, by proposing a method which allows a simple and robust regulation of the operation of a rocket engine comprising at least one combustion chamber and a first liquid propellant feed circuit with a first pump for a first liquid propellant and a first turbine for actuating the first pump, a first supply valve and a control device of the first turbine.
    In at least one embodiment, this object is achieved by virtue of the fact that this regulation method comprises the following steps: calculation of an opening set point of the first supply valve, from an external setpoint, following an open loop control law; calculating a setpoint for the control device of the first turbine, from said external setpoint and at least one feedback value, according to a closed-loop control law; and controlling the first supply valve and the control device of the first turbine, according to the respective instructions.
    Thanks to these arrangements, the closed-loop control of the first turbine makes it possible to correct the errors due to the open-loop control of the first supply valve, thus permitting precise regulation of the rocket engine even with relatively high computing power. low and acting on a limited number of regulating devices.
    This control method is particularly applicable to a rocket engine with an "expander" cycle, in which said first supply circuit comprises a heat exchanger for heating the first liquid propellant downstream of the first pump, and passes through the first turbine downstream of the heat exchanger to actuate the first pump by expansion in the first turbine, the first liquid propellant heated in the heat exchanger. However, it is also applicable to other types of rocket engines, such as for example rocket engines with a gas generator, comprising a gas generator for supplying at least the first turbine with hot combustion gases. In this case, the control device of the first turbine may include in particular at least one supply valve of the gas generator. Furthermore, "gas generator" must be understood in a broad sense encompassing all total or partial combustion devices that can generate hot combustion gases, including a prechamber of combustion in a staged combustion rocket engine.
    Regardless of whether the rocket engine is an "expander" rocket engine, a gas-powered rocket engine or a rocket engine of another type, the control device of the first turbine may include a bypass valve of the first turbine. Such a bypass valve allows a regulation of the flow of engine fluid passing through the turbine and therefore, indirectly, the flow rate of the first propellant pumped by the first pump and thus provided to the combustion chamber through the first supply valve.
    Said external setpoint may in particular comprise a set value of gas pressure in the combustion chamber. In this case, the at least one feedback value can then comprise a measured value of gas pressure in the combustion chamber, so that the closed-loop control law corrects an error between the set value and the measured value. . However, the use of an estimated value, rather than a measured value, as a feedback value is also conceivable.
    In order to limit the stresses exerted on the first supply valve and / or the control device of the first turbine, the method may further comprise a step of filtering the external setpoint in a tracking filter, and the calculation of the opening setpoint of the first fuel supply valve of the combustion chamber is then performed from the filtered external setpoint in this filtering step, and / or the calculation of the setpoint for the control device of the first turbine is performed by a disturbance corrector from the at least one feedback value and the external setpoint filtered in this filtering step, the disturbance corrector having a cutoff frequency higher than the tracking filter.
    Although the regulation method can be used for the regulation of a single-engined or hybrid rocket engine, the combustion chamber of which is thus fed with only one liquid propellant, it is also applicable to a rocket engine further comprising a second liquid propellant supply circuit for a second liquid propellant. Said external setpoint may then comprise a setpoint ratio value between a flow rate of the first propellant and a flow rate of the second propellant. In this case, said at least one feedback value may comprise a measured value of the ratio between the flow rate of the first propellant and the flow rate of the second propellant, which can then be compared to its setpoint in the closed loop of regulation of the first propellant. turbine.
    The second liquid propellant feed circuit may in particular comprise a second feed valve, and the control method also comprises a step of calculating an opening set point of the second feed valve, starting from external setpoint, following an open loop control law, and the control of the second supply valve following this opening instruction.
    Furthermore, although other means of ensuring the circulation of the second propellant are also conceivable, such as a pressurized tank, the second liquid propellant supply circuit may include a second pump for said second propellant liquid. This second pump could be actuated by the first turbine or by other means conceivable by the person skilled in the art, but a particularly feasible option is that the rocket engine comprises a second turbine for actuating the second pump. The rocket engine could then also include a device for regulating the second turbine, and the control method also comprises calculating a setpoint for the control device of the second turbine, from said external setpoint and from minus a feedback value, according to a closed-loop control law, and the control of the second turbine control device, according to this instruction for the second turbine control device. The circulation of the two propellants would therefore be regulated in a similar manner. The control device of the first turbine and that of the second turbine may also form only one and / or share some elements.
    The present disclosure also relates to a control circuit of a rocket motor with at least one liquid propellant, said rocket engine comprising at least one combustion chamber and a first liquid propellant feed circuit with a first pump for a first propellant. liquid and a first turbine for actuating the first pump, a first supply valve and a regulating device of the first turbine. In at least one embodiment, this control circuit comprises an open control loop of at least the first supply valve, with a module for calculating an opening setpoint of the first supply valve of the combustion chamber from an external setpoint, and a closed control loop of the first turbine, with a calculation module of a setpoint for the control device of the first turbine from said external setpoint and from minus a feedback value.
    Furthermore, this disclosure also relates to a rocket engine equipped with such a control circuit and software and / or computer support containing instructions for executing the control method by a programmable electronic control unit. By "computer support" is meant any medium for storing data, in a sustainable manner and / or transient, and their subsequent reading by a computer system. Thus, "computer medium" is understood to mean, inter alia, magnetic tapes, magnetic and / or optical disks, or solid state electronic memories, both volatile and non-volatile.
    BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood and its advantages will appear better on reading the following detailed description of embodiments shown by way of non-limiting examples. The description refers to the accompanying drawings in which: - Figure 1 is a schematic illustration of a biergol rocket engine "expander" cycle according to one embodiment; FIG. 2 is a block diagram of a rocket motor control circuit of FIG. 1; FIG. 3 is a schematic illustration of a staged combustion biegol rocket engine; and FIG. 4 is a block diagram of a rocket motor control circuit of FIG. 3.
    Detailed description of the invention
    FIG. 1 schematically illustrates an "expander" rocket engine 1, such as can be incorporated, for example, in a space launcher stage, an orbital transfer vehicle or another spacecraft.
    This rocket engine 1 comprises a propulsion chamber 2, a first supply circuit 10, a second supply circuit 20, and an electronic control unit 50. The propulsion chamber 2 comprises a combustion chamber 3, an igniter 5 and a nozzle 4 for the expansion and supersonic ejection of high enthalpy gas generated by the combustion of a mixture of a first and a second propellant liquid in the combustion chamber 3. The first supply circuit 10 connects a first reservoir 11 , containing the first liquid propellant, to the propellant chamber 2, for supplying the combustion chamber 3 with this first liquid propellant. The second supply circuit 20 connects a second tank 21, containing the second liquid propellant, to the propellant chamber 2, to supply the combustion chamber 3 with the second liquid propellant.
    In the illustrated embodiment, the first supply circuit 10 comprises, successively in the direction of flow of the first propellant of the first reservoir 11 inside the combustion chamber 3, a first pump 12, a heat exchanger 13, a first turbine 14, a second turbine 24, and a first valve 15 for supplying the combustion chamber 3. The second supply circuit 20 comprises, successively in the direction of flow of the second propellant of the second reservoir 21 inside the combustion chamber 3, a second pump 22 and a second valve 25 for supplying the combustion chamber 3. In addition, first and second flow sensors 16, 26 are installed on, respectively, the first and second feed circuit 10, 20 to measure the flow (volume or mass) of each propellant supplied to the combustion chamber. In the illustrated embodiment, these flow sensors 16, 26 are situated directly upstream of the valves 15, 25 for supplying the combustion chamber, but other alternative positions are also conceivable, as well as other means for obtaining this magnitude such as an estimator or a model. In addition, a pressure sensor 30 is installed inside the combustion chamber 3. The heat exchanger 13 is formed on at least one wall of the propulsion chamber 2, so as to heat the first propellant with heat. combustion propellant in the combustion chamber 3, while ensuring the cooling of this wall. The first turbine 14 is mechanically coupled to the first pump 12, thus forming a first TPI turbopump. On the other hand, the second turbine 24 is mechanically coupled to the second pump 22, thus forming a second turbopump TP2. Thus, a first partial expansion of the first propellant, after having been heated in the heat exchanger 13, in the first turbine 14, can actuate the first pump 12 to supply the combustion chamber 3 with the first propellant, while a second partial expansion of the first propellant can then actuate the second pump 22 to supply the combustion chamber 3 with the second propellant. Furthermore, the first power supply circuit 10 also comprises a branch 17 for bypassing the first and second turbines 14, 24, with a 17v valve bypassing the first and second turbines 14, 24, as well as a branch 18 for bypassing the second turbine 24, with a valve 18v bypassing the second turbine 24. The bypass valves 17v, 18v thus form regulating devices of the first and second turbines 14,24.
    In the illustrated embodiment, flow rate sensors 16, 26 and 30 as well as feed valves 15, 25 and bypass 17v, 18v are all connected to control electronics 50 to form a circuit. 60 of the rocket motor 1, shown in Figure 2. This electronic control unit 50 may include a digital computer with a predetermined sampling rate.
    As illustrated in FIG. 2, this regulation circuit 60 can comprise a tracking filter 70 upstream of an open loop 80 for controlling the supply valves 15, 25 and a closed loop 90 for controlling the turbines 14. 24. The tracking filter 70 can in particular be a first or second order tracking filter. Although in the illustrated embodiment this filter is common to the two loops 80, 90, it is also possible to have a tracking filter specific to each loop. This filter is intended to filter an external setpoint Cext comprising setpoints pgc, c. _ext St Tc_ext for, respectively, a gas pressure in the combustion chamber 3 and a flow ratio between the two propellants, thus resulting in a filtered external setpoint Cmt with filtered set values Pgc, c_fiit and rc_fiit. The tracking filter 70 can in particular be a first-order digital filter of the formula Cmt (k) = KmiCmtCk-1) + Kfiit2Cext (kl), in which each of the filtered setpoint values Pgc, c_fiit and rc_fiit at a time of sampling k corresponds to the addition of the filtered setpoint value gc, c_fiit and rcmt corresponding to the previous sampling time k-1, multiplied by a first predetermined constant Kniti, and the corresponding setpoint value pgc, c_ext and rc_ext / at the moment k-1 also, multiplied by a second predetermined constant K r, it2- For example, for a motor whose response is desired in minimum 3s, with a sampling rate of 10 ms, the value of K ^ i can be 0.99005 and that of Κηκ 0.00995.
    The open loop 80 comprises a first calculation module 81 connected to the tracking filter 70 for calculating DVCO instructions, DVCH opening of the supply valves 15,25 of the combustion chamber 3 from the filtered external setpoint Cnit. This calculation module 81 can in particular be configured to apply a polynomial control law, such as that according to the following equations: DVCO = KiX4 + K2x3 + K3x2 + K4X + K5 DVCH = K6X4 + K7x3 + K8X2-i-K9X + Kio in which Ki to Km are predetermined coefficients and x is a ratio between the gas pressure setpoint filtered value gc, c_fiit and a predetermined maximum value gc, max for this gas pressure in the combustion chamber.
    However, other types of control law, such as an artificial neural network control law, are also conceivable. The calculation module 81 is connected to actuators of the supply valves 15, 25 of the combustion chamber 3 to transmit them the opening instructions DVCO, DVCH which correspond to a ratio between the displacement of each actuator from a closed position of the corresponding valve, and its total travel between this closed position and a maximum open position of the valve.
    The closed loop 90 comprises a divider module 92, possibly protected against divisions by 0, and a calculation module 91 connected to the tracking filter 70, to the pressure sensor 30 and, through the divider module 92, to the flow sensors 16 , 26 for calculating DVBPH setpoints, DVBPO for opening the bypass valves 17v, 18v from the filtered external setpoint Cmt, and feedback values comprising a gas pressure value pgc, m measured by the pressure sensor 30 , and a ratio rm between a flow rate Qi measured by the flow sensor 16 and a flow rate Q2 measured by the flow sensor 26. The calculation module 91, which has a higher cut-off frequency than the tracking filter 70, can for example, take the form of a dual integral proportional disturbance corrector like that illustrated in FIG. 2, with a first integral proportional corrector 93 for generating the DVBPH setpoint for opening the gate valve. 17v from the error errpgC between the filtered gas pressure setpoint value gcc, c_fiit and the gas pressure value gcg, m measured by the pressure sensor 30, and a second integral proportional corrector 94 to generate the DVBPO setpoint for opening the bypass valve 18v from the error errr between the filtered value of the ratio of propellant flow rates rc_nit and the ratio rm between the flow rate Qi measured by the flow sensor 16 and the flow rate Q2 measured by the flow sensor 26, according to the formulas: DVBPH (k) = DVBPH (kl) -Kpierrpgc (k) + Kpierrpgc (kl) -KiiTeerrpgc (kl) DVBPO (k) = DVBPO (k1) -Kp2errr (k) ) + Kp2errr (kl) -Ki2Teerrr (kl) in which DVBPH (k) and DVBPO (k) will be the values of the DVBPH and DVBPO setpoints, respectively, at the sampling time k, DVBPH (kl) and DVBPO (kl ) the values of these same setpoints at the previous sampling time k-1, errpgc (k) and errr (k) the error values errpgc and errr, respectively, at the sampling instant k, errpgC (kl) and errr (kl) the values of these same errors at the previous sampling instant, Kpi and Kp2 the proportional constants of, respectively, the first and the second integral proportional corrector 93 and 94, Ku and Ki2 their respective integral constants and Te the sampling rate of the calculation module 91. The proportional constants Kpi and Kp2 may for example have respective values of 8.1 and 90 MPa'1 and the integral constants Κ, ι and Ki2 of the respective values of 54 and 6000 MPa'V1 with a sampling rate Te of 10 ms.
    However, this computation module 91 may alternatively take other forms, such as for example that of a multivariable corrector, and in particular a predictive internal model corrector.
    In operation, the first liquid propellant, which may especially be a cryogenic fluid such as liquid hydrogen, is extracted from the first reservoir 11 and pumped by the first pump 12, through the first supply circuit 10, to the chamber of In the heat exchanger 13, the first liquid propellant is heated, thereby increasing its enthalpy. Part of this additional enthalpy is then used to actuate the first pump 12 and the second pump 22 by partial expansion of at least a portion of the flow of the first liquid propellant in, respectively, the first turbine 14 and the second turbine 24. The conduits bypass 17, 18, with their respective bypass valves 17v, 18v, allow the regulation of the proportion of the total flow of the first liquid propellant which passes through each of these two turbines 14,24 and thus that of the regime of the two turbopumps TPI, TP2 . The total flow rate of the first liquid propellant then passes through the first supply valve 15 of the combustion chamber 3, which contributes to regulating this total flow rate, before being injected into the combustion chamber 3. The second liquid propellant, which can also be a cryogenic fluid such as liquid oxygen, is extracted from the second tank 21 and pumped by the second pump 22, through the second supply circuit 20, including the second valve 25 supplying the combustion chamber 3, to be injected into the combustion chamber 3. The second feed valve 25 contributes to regulate its flow.
    The two propellants, mixed within the combustion chamber 3, are in combustion, with a large amount of heat, to generate high enthalpy combustion gases whose expansion and acceleration to a supersonic velocity in the nozzle 4 produces a thrust in the opposite direction. At the same time, a part of the heat released by the combustion of the propellants contributes to heating the first liquid propellant circulating through the heat exchanger 13.
    The thrust thus produced by the rocket engine 1 depends in particular on the pressure of the gas pgc, m in the combustion chamber, as well as their temperature, and thus indirectly the flow rates Qi, Q2 of the two propellants. These values are measured by the flow rate sensors 16, 26 and 30 and transmitted to the electronic control unit 50.
    In the electronic control unit 50, the tracking filter 70 filters the external setpoint Cext comprising setpoints pgc, c_ext and rc_ext and transmits the filtered external setpoint Cnit to the open loop 80 for controlling the supply valves 15, 25 and the closed loop 90 for controlling the turbines 14,24. This external instruction Cext can for example follow a preprogrammed profile stored in an internal memory of the electronic control unit 50 to follow a preprogrammed profile, to be calculated in flight by the electronic control unit 50 as a function of flight data of the powered by the rocket engine 1, or be transmitted to this machine from a base station.
    In the open control loop of the supply valves 15, 25, the first calculation module 81 connected to the tracking filter 70 calculates instructions DVCO, DVCH opening of the supply valves 15, 25 of the combustion chamber 3 from the filtered external setpoint Cnit. These opening instructions DVCO, DVCH are then transmitted to the actuators of the supply valves 15, 25 for controlling the supply of the combustion chamber 3 with propellants, and therefore the operating speed of the rocket engine 1.
    The closed loop 90 for controlling the turbines 14,24 makes it possible to obtain a more precise regulation of this operating regime. In this closed loop, the filtered setpoint values pgc, fiit and rfl | t of the filtered external setpoint Cnit are compared with the respective measured values pgc, m and rm to obtain the corresponding errors errpgc and errr, from which the calculation module 91 calculates the DVBPH, DVBPO setpoints for opening bypass valves 17v, 18v. These opening instructions DVBPH, DVBPO are then transmitted to the actuators of the bypass valves 17v, 18v to regulate the speeds of the turbines 14,24, and correct the start of engine speed rocket relative to the external setpoint Ce *. The electronic control unit 50, and in particular the calculation module 81 of the open loop 80 can be configured such that the DVBPH, DVBPO opening setpoints of the bypass valves 17v, 18v in the closed loop 90 remain in a range between 20 and 80% of opening. In particular, this can be achieved by appropriate choice of coefficients Ki to Ki0.
    Although in this first embodiment, the method and control system serve to regulate the operation of an "expander" rocket motor, it is also conceivable to adapt them to the regulation of other types of rocket motors , such as the staged combustion rocket motor illustrated in FIG.
    This rocket engine 1 also comprises a propulsion chamber 2, a first supply circuit 10, a second supply circuit 20, and an electronic control unit 50. The propulsion chamber 2 comprises a combustion chamber 3, an igniter 5 and a nozzle 4 for the expansion and supersonic ejection of high enthalpy gas generated by combustion in the combustion chamber 3. The first supply circuit 10 connects a first reservoir 11, containing the first liquid propellant, to a gas generator 100 forming prechamber of combustion, for supplying this gas generator 100 in this first liquid propellant. The second supply circuit 20 connects a second tank 21, containing the second liquid propellant, to the propellant chamber 2 and the gas generator 100, to supply the combustion chamber 3 and the gas generator 100 in the second liquid propellant. Finally, conduits 101, 102 connect the gas generator 100 to the propulsion chamber 2 to supply the combustion chamber with the gases resulting from a partial combustion, in the gas generator 100, of a mixture of the said first and second propellants, a mixture rich in this first propellant.
    In this second embodiment, the first supply circuit 10 comprises, successively in the direction of flow of the first propellant, a first pump 12, a first supply valve 15 and a heat exchanger 13. The first circuit supply 10 further comprises, for driving the first pump 12, a first turbine 14, through which the conduit 101 passes. The second supply circuit 20 comprises, successively in the direction of flow of the second propellant, a second pump 22 and a second supply valve 25. It also includes, for driving the second pump 22, a second turbine 24, through which the conduit 102 passes, and a bypass 103, with a valve 104, for supplying the gas generator 100. Moreover, first and second flow sensors 16, 26 are installed on, respectively, the first and second feed circuits 10, 20 to measure the flow (volume or mass) of each propellant supplied to the combustion chamber. In this second embodiment, these flow sensors 16, 26 are located directly downstream of the pumps 12, 22, but other alternative positions are also possible as well as other means for obtaining this magnitude such as an estimator or A model. In addition, a pressure sensor 30 is installed inside the combustion chamber 3. The heat exchanger 13 is formed on at least one wall of the propulsion chamber 2, so as to heat the first propellant with heat. combustion propellant in the combustion chamber 3, while ensuring the cooling of this wall. The first turbine 14 is mechanically coupled to the first pump 12, thus forming a first TPI turbopump. On the other hand, the second turbine 24 is mechanically coupled to the second pump 22, thus forming a second turbopump TP2. Thus, partial combustion, in the gas generator 100, of the first propellant with a portion of the second propellant, derived from a main flow of this second propellant between the second pump 22 and the second valve 25, generates a rich gas mixture. in this first propellant and a first portion, flowing from the gas generator 100 to the propellant chamber 2 through the first duct 101, can actuate the first pump 12 by partial expansion in the first turbine 14 to ensure the circulation of the first propellant, while a second part, flowing in parallel from the gas generator 100 to the propulsion chamber 2 through the second conduit 102, can actuate the second pump 22 by partial expansion in the second turbine 24 to ensure the circulation of the second propellant. Furthermore, the assembly also comprises a valve 105 installed on the conduit 102. The valves 104 and 105 thus form a device for regulating the turbines 14 and 24.
    In this second embodiment, the flow rate sensors 16, 26 and 30 as well as the valves 15, 25, 104 and 105 are all connected to the control electronics unit 50 to form the motor control circuit 60. FIG. 4. As in the first embodiment, the electronic control unit 50 of this second embodiment may also comprise a digital computer with a predetermined sampling rate.
    The control circuit 60 according to this second embodiment is largely similar to that of the first embodiment, and differs from it mainly in that its closed loop 90 regulates the operation of the turbines 14,24 through the valves 104,105 , rather than through bypass valves. For this, the calculation module 91 calculates instructions DPBOV, DHGV opening the valves 104, 105 similarly to the instructions DVBPH, DVBPO of the first embodiment. The other elements in this second embodiment are equivalent to those of the first embodiment and receive the same references in the drawing.
    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. In addition, individual features of the various embodiments mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.
    权利要求:
    Claims (15)
    [1" id="c-fr-0001]
    A method of controlling a rocket motor (1) with at least one liquid propellant, said rocket engine comprising at least one combustion chamber (3) and a first liquid propellant supply circuit (10) with a liquid propellant. first pump (12) of a first liquid propellant and a first turbine (14) for the actuation of the first pump (12), a first supply valve (15) and a control device of the first turbine (14). ), this control method comprising the following steps: calculation of an opening setpoint (DVCH) of the first supply valve (15), from an external setpoint (Cext), according to a control law in open loop; calculating a setpoint for the control device of the first turbine (14), from said external setpoint (Cext) and at least one feedback value, according to a closed-loop control law; and controlling the first supply valve (15) of the combustion chamber (3) and the control device of the first turbine (14) according to the respective instructions.
    [2" id="c-fr-0002]
    2. A method of controlling a rocket motor (1) according to claim 1, wherein said first supply circuit (10) comprises a heat exchanger (13) for heating the first liquid propellant downstream of the first pump (12), and passes through the first turbine (14) downstream of the heat exchanger (13) to actuate the first pump (12) by expansion in the first turbine (14) of the first liquid propellant heated in the heat exchanger (13).
    [3" id="c-fr-0003]
    The method of regulating a rocket engine (1) according to claim 1, wherein the rocket engine (1) further comprises a gas generator (100) for supplying at least the first turbine (14) with gas hot combustion.
    [4" id="c-fr-0004]
    4. A method of regulating a rocket engine (1) according to claim 3, wherein the regulating device of the first turbine (14) comprises at least one valve (104) for supplying the gas generator (100). .
    [5" id="c-fr-0005]
    A method of controlling a rocket motor (1) according to any one of the preceding claims, wherein the regulating device of the first turbine (14) comprises a bypass valve (17v) of the first turbine (14). ).
    [6" id="c-fr-0006]
    6. A method of regulating a rocket engine (1) according to any one of the preceding claims, wherein said external setpoint (Ce *) comprises a gas pressure reference value (Pgc_ext) in the combustion chamber ( 3).
    [7" id="c-fr-0007]
    The method of regulating a rocket motor (1) according to claim 6, wherein the at least one feedback value comprises a measured value of gas pressure (pgc_m) in the combustion chamber (3).
    [8" id="c-fr-0008]
    8. A method of regulating a rocket motor (1) according to any one of the preceding claims, comprising a step of filtering the external setpoint (Cext) in a tracking filter (70) to obtain a filtered external setpoint ( Cmt), and wherein the calculation of the opening setpoint (DVCH) of the first supply valve (15) is made from the filtered external setpoint (Cnit) and / or the calculation of the setpoint for the device for regulating the first turbine (14) is performed by a disturbance corrector from the at least one feedback value and the filtered external setpoint (Cnit), the disturbance corrector having a higher cut-off frequency than the tracking filter.
    [9" id="c-fr-0009]
    9. A method of controlling a rocket motor (1) according to any one of the preceding claims, wherein the rocket motor (1) further comprises a second feed circuit (20) for a second liquid propellant.
    [10" id="c-fr-0010]
    10. Control method according to claim 9, wherein said external setpoint (Cext) comprises a set value (rext) ratio between a flow (Qi) of the first propellant and a flow (Q2) of the second propellant.
    [11" id="c-fr-0011]
    11. Control method according to claim 10, wherein said at least one feedback value comprises a measured value (rm) of the ratio between the flow rate (Qi) of the first propellant and the flow rate (Q2) of the second propellant.
    [12" id="c-fr-0012]
    A method of regulating a rocket motor according to any of claims 9 to 11, wherein the second liquid propellant feed circuit (20) comprises a second feed valve (25), and the method control device also comprises a step of calculating an opening setpoint (DVCO) of the second supply valve (25), from the external setpoint (Cext), according to an open-loop control law, and the control of the second supply valve (25) following this opening instruction (DVCO).
    [13" id="c-fr-0013]
    13. A method of regulating a rocket motor (1) according to any one of claims 9 to 12, wherein the second liquid propellant supply circuit (20) comprises a second pump (22) for said second propellant liquid.
    [14" id="c-fr-0014]
    14. A method of regulating a rocket motor (1) according to claim 13, wherein the rocket motor (1) comprises a second turbine (24) for actuating the second pump (22) and a device regulating the second turbine (24), and the control method also comprises calculating a setpoint for the control device of the second turbine (24), from said external setpoint (Cext) and at least one feedback value, according to a closed loop control law, and the control of the second turbine control device (24), according to this setpoint for the second turbine control device (24).
    [15" id="c-fr-0015]
    15. Control circuit of a rocket motor (1) to at least one liquid propellant, said rocket engine (1) comprising at least one combustion chamber (3) and a first supply circuit (10) liquid propellant with a first pump (12) a first liquid propellant and a first turbine (14) for the actuation of the first pump (12), a first supply valve (15) and a control device of the first turbine (14). ), the control circuit comprising: an open loop (80) for controlling the first supply valve (15), with a module (81) for calculating at least one opening setpoint (DVCH) of the first supply valve (15) from an external setpoint (Cext); and a closed loop (90) for controlling the first turbine (14), with a module (91) for calculating a setpoint for the device for regulating the first turbine (14) from said external setpoint (Cext) and at least one feedback value.
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    公开号 | 公开日 | 专利标题
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    同族专利:
    公开号 | 公开日
    JP6254238B2|2017-12-27|
    EP3153691A1|2017-04-12|
    EP3153691B1|2020-06-03|
    US20170101963A1|2017-04-13|
    FR3042227B1|2020-04-03|
    RU2016139441A|2018-04-09|
    JP2017072138A|2017-04-13|
    引用文献:
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    US3851157A|1973-07-09|1974-11-26|United Aircraft Corp|Self-correcting feedback control system|
    US7477966B1|2004-02-20|2009-01-13|Lockheed Martin Corporation|Propellant management system and method for multiple booster rockets|
    FR2928464A1|2008-03-07|2009-09-11|Snecma Sa|CONTROL DEVICE FOR A REGULATED AND ENGINE SYSTEM INCORPORATING SUCH A DEVICE|EP3581782A1|2018-06-13|2019-12-18|ArianeGroup SAS|Rocket engine control method and control device|
    RU2742516C1|2019-08-08|2021-02-08|Александр Вячеславович Дыбой|Propulsion system with a rocket engine|US3737103A|1969-08-28|1973-06-05|Trw Inc|Digital liquid vector control system|
    JPH10238408A|1997-02-21|1998-09-08|Ishikawajima Harima Heavy Ind Co Ltd|Expander cycle engine|
    US7900436B2|2007-07-20|2011-03-08|The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration|Gas-generator augmented expander cycle rocket engine|
    FR2937092B1|2008-10-15|2010-12-10|Snecma|METHOD AND DEVICE FOR CALCULATING A SEQUENCE OF STARTING OR STOPPING AN ENGINE|
    US9677504B2|2010-05-11|2017-06-13|Orbital Atk, Inc.|Rockets, methods of rocket control and methods of rocket evaluation utilizing pressure compensation|
    US8250853B1|2011-02-16|2012-08-28|Florida Turbine Technologies, Inc.|Hybrid expander cycle rocket engine|
    FR2997731B1|2012-11-06|2018-07-27|Arianegroup Sas|DEVICE AND METHOD FOR SUPPLYING A ROCKER ENGINE|US11181076B2|2016-03-03|2021-11-23|Kevin Michael Weldon|Rocket engine bipropellant supply system including an electrolyzer|
    KR101942001B1|2017-08-11|2019-01-25|충남대학교산학협력단|Frequency variable type burner|
    CN112502856B|2020-11-25|2021-11-19|北京航空航天大学|Liquid oxygen and oxygen double-path adjustable supply system|
    法律状态:
    2016-10-13| PLFP| Fee payment|Year of fee payment: 2 |
    2017-04-14| PLSC| Search report ready|Effective date: 20170414 |
    2017-10-23| PLFP| Fee payment|Year of fee payment: 3 |
    2018-10-23| PLFP| Fee payment|Year of fee payment: 4 |
    2019-10-23| PLFP| Fee payment|Year of fee payment: 5 |
    2020-10-22| PLFP| Fee payment|Year of fee payment: 6 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1502103|2015-10-08|
    FR1502103A|FR3042227B1|2015-10-08|2015-10-08|METHOD AND CIRCUIT FOR MOTOR ROCKET CONTROL|FR1502103A| FR3042227B1|2015-10-08|2015-10-08|METHOD AND CIRCUIT FOR MOTOR ROCKET CONTROL|
    JP2016199222A| JP6254238B2|2015-10-08|2016-10-07|Method and circuit for tuning a rocket engine|
    US15/288,521| US20170101963A1|2015-10-08|2016-10-07|Method and a circuit for regulating a rocket engine|
    RU2016139441A| RU2016139441A|2015-10-08|2016-10-07|METHOD AND CONTOURS OF ROCKET ENGINE REGULATION|
    EP16192935.1A| EP3153691B1|2015-10-08|2016-10-07|Rocket engine control circuit and method|
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