• 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 space propulsion and more particularly of electric space propulsion. A space propulsion system (100) according to the invention comprises at least one electrostatic propellant (101) with at least a first electrical consumer, a resistor (102), a propellant supply circuit (104), and a circuit Power supply unit (103) comprising at least a first power supply line (131) and a first switch (114-1, 114'-1,114 "-1) for selecting between connecting said first power supply line (131). to the resistor (102) and connecting said first power supply line (131) to said first electrical consumer of the electrostatic thruster (101), thereby enabling the application of a space propulsion method including a switching step for selecting a first propulsion mode, wherein the resistor (102) is activated, or a second propulsion mode, wherein the electrostatic propellant (101) is activated.
    公开号:FR3024436A1
    申请号:FR1457371
    申请日:2014-07-30
    公开日:2016-02-05
    发明作者:Frederic Marchandise
    申请人:SNECMA SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The present invention relates to the field of space propulsion. In this field, electric thrusters are becoming more and more common, especially for the control of orbit and orbit of spacecraft. Indeed, the different types of electric thrusters available offer a specific pulse generally higher than conventional chemical or cold gas thrusters, thus reducing the propellant fluid consumption for the same maneuvers, with consequent increase in the duration of life and / or the payload of spacecraft. Among the different types of electric thrusters, two categories are known in particular: so-called thermoelectric thrusters, in which the propellant fluid is electrically heated prior to its expansion in a propellant nozzle, and those called electrostatic, in which the propellant fluid is ionized and directly accelerated by an electric field. Amongst the thermoelectric thrusters, in particular those referred to as resistors, in which the heat is transmitted to the propellant fluid by at least one resistance heated by the Joule effect. On the other hand, among the electrostatic thrusters, the so-called Hall effect thrusters are included. In these thrusters, also known as a closed-electron drift plasma engine or stationary plasma engines, electrons emitted by a transmitting cathode are captured by a magnetic field generated by coils located around and in the center of a discharge channel. annular section, thus forming a virtual cathode grid at the end of this discharge channel. The propellant fluid (typically xenon in the gaseous state) is injected at the bottom of the discharge channel and electrons escaping from this virtual cathode gate to an anode located at the bottom of the discharge channel impel 30 molecules of the propellant fluid, ionizing this one, which is consequently accelerated towards the virtual cathode grid by the electric field reigning between this one and the anode, before being neutralized by other electrons emitted by the emitting cathode. Typically, in order to ensure the emission of electrons from the cathode, the cathode is electrically heated.
    [0002] Hall effect thrusters are also not the only ones comprising similar emitting cathodes. Another example of an electrostatic thruster with a similar cathode is the so-called high efficiency multi-stage plasma thruster or HEMP, according to its acronym (High Efficiency Multistage Plasma Thruster), described for example by H.-P. Harmann, N. Koch and G. Kornfeld in "Low Complexity and Low Cost Electric Propulsion System for Telecom Satellites Based on HEMP Thruster Assembly", IEPC-2007-114, 30th International Electric Propulsion Conference, Florence, Italy, 17-20 September 2007 .
    [0003] In such a HEMP propellant, the ionized propellant fluid is accelerated by an electric field formed between an anode and a plurality of virtual cathode grids formed by electrons trapped by the magnetic fields of a plurality of permanent magnets. In general, all the electrostatic thrusters comprise an emitter cathode at least to neutralize the propellant fluid downstream of the propellant. Electrostatic thrusters provide particularly high specific pulses in comparison with other types of thrusters, including thermoelectric thrusters. On the other hand, their thrust is very low. Space propulsion systems have therefore been proposed combining electrostatic thrusters for slow maneuvers, such as the maintenance of orbit or the desaturation of reaction wheels, and propellants of other types for maneuvers requiring higher thrusts. Thus, Mr. De Tata, P.-E. Frigot, S. Beekmans, H. Lübberstedt, D. Birreck, A. Demairé and P.
    [0004] Rathsman, in "SGEO Development Status and Opportunities for the EP-based Small European Telecommunication Platform", IEPC-2011-203, 32nd International Electric Propulsion Conference, Wiesbaden, Germany, 11-15 September 2011 and S. Naclerio, J. Soto Salvador , E. Such, R. Avenzuela and R. Perez Vara, in "Small GEO Xenon Propelling Supply Assembly Pressure Regulator Panel: Test Results and Comparison with ECOSIMPRO Predictions", SP2012-2355255, 3rd International Conference on Space Propulsion, Bordeaux, 7- May 10, 2012, have described a space propulsion system for small geostationary satellites, comprising electrostatic thrusters and cold gas thrusters powered by a common propellant fluid supply circuit. However, the specific impulse of cold gas thrusters is very limited, which makes them very greedy propellant fluid for maneuvers high thrust, and also the pooling of resources between the different types of thrusters in this system is not very important, resulting in a fairly high complexity of the system.
    [0005] OBJECT AND SUMMARY OF THE INVENTION The present invention aims to remedy these drawbacks. In particular, this disclosure aims at proposing a space propulsion system that makes it possible to offer at least a first propulsion mode, with a high specific impulse but a low thrust, and a second mode of propulsion, with a higher thrust and a smaller specific impulse. that of the first mode, but however higher than that can be provided by cold gas thrusters, and this with a relatively simple power supply circuit. In at least one embodiment, this object is achieved by virtue of the fact that the propulsion system comprises an electrostatic propellant with at least a first electrical consumer, a resistor, a propellant supply circuit, and a power supply circuit. electrical connector having at least a first power supply line and a first switch for selecting between connecting said first power supply line to the resistor and connecting said first power supply line to said first electrical consumer of the electrostatic propellant. The use of a resist-jet makes it possible to obtain a higher specific impulse than the cold-gas thrusters, while making it possible to continue to share at least a portion of the propellant fluid supply circuit for the propellant fluid supply of the propellant. electrostatic thruster and resistor. At the same time, the first switch makes it possible to supply the resistojet with power from the same power supply line serving alternately to supply electricity to a first electrical consumer of the electrostatic thruster, which simplifies the circuit of power supply. In particular, said first electrical consumer of the electrostatic propellant may comprise a heating element of a cathode emitting said electrostatic thruster. Since the heating elements of such emitting cathodes and those of the resistors are normally formed by similar electrical resistances, it is particularly easy to supply them electrically from the same source without conversion or transformation of current or voltage.
    [0006] In order to control the supply of the electrostatic propellant and the propellant fluid resistor, said propellant fluid supply circuit comprises at least one electrostatic propellant supply valve and at least one supply valve of the resistor. In particular, the propulsion system may further comprise at least one valve opening control line and a second switch for selecting between connecting said valve opening control line to the electrostatic propellant supply valve and connect said valve opening control line to the at least one supply valve of the resistor. Depending on the selected propulsion mode, the same valve opening control line can thus alternately control the supply of the electrostatic propellant or the propulsion fluid resistojet, thus simplifying the control of the valves. Typically, in the electrostatic thrusters, a particularly high electrical voltage must be established between a cathode and an anode in order to generate the electric field ensuring the acceleration of the ionized propulsive fluid. This voltage is normally substantially higher than that of the power supply of the heating of the emitting cathode, or that supplied by the electrical sources on board a spacecraft, such as photovoltaic panels, batteries, fuel cells or thermoelectric generators. To also provide this high voltage, said power supply circuit may further comprise at least one electrical conditioning unit capable of supplying at least one other electrical consumer of the electrostatic thruster at a voltage substantially higher than the first electrical consumer. The first power supply line may be at least partially integrated into said electrical conditioning unit, although it may also alternately bypass the electrical conditioning unit and be connected directly to a distribution bar of a machine electrical network space or on-board power sources.
    [0007] Said power supply circuit may comprise at least one thruster selection unit in which at least said first switch is integrated. Thus, if several connections must be switched simultaneously to select one or the other thruster, all the corresponding switches may optionally be integrated in such a thruster selection unit, and be controlled by the same command. Said electrostatic propellant may in particular be a Hall effect thruster. Indeed, the Hall effect thrusters have already proven their reliability in space propulsion. However, other types of electrostatic propellant are also conceivable, including HEMP thrusters. In particular to ensure propulsion on several different axes, this space propulsion system may comprise a plurality of electrostatic thrusters. In this case, in order to simplify the supply of propellant gas to this assembly, the propellant fluid supply circuit may comprise at least one pressure regulating device common to several of said electrostatic thrusters. However, complementarily or alternatively to at least one pressure regulating device common to several of said electrostatic thrusters, a propellant fluid supply circuit may comprise an individual pressure regulating device for at least one of said electrostatic thrusters. The present disclosure also relates to an orientation and / or trajectory control system comprising such a space propulsion system, a spacecraft, for example a satellite or a probe, comprising such a space propulsion system, as well as a method space propulsion system comprising a step of switching between an electrostatic thruster and a resistor, wherein a first switch is used to connect a first power supply line to the resistor or a first low-voltage electrical consumer of the electrostatic thruster, to select a first propulsion mode, in which the resistojet is activated, or a second propulsion mode, in which the electrostatic thruster is activated.
    [0008] BRIEF DESCRIPTION OF THE DRAWINGS The invention will be better understood and its advantages will appear better on reading the detailed description which follows, of several embodiments shown by way of non-limiting examples. The description refers to the accompanying drawings in which: - Figure 1 is a schematic view of a spacecraft equipped with an orientation control system and trajectory with a space propulsion system according to any one of the embodiments ; FIG. 2A is a detailed diagram illustrating a space propulsion system according to a first embodiment, with switches in the position of selection of an electrostatic thruster; FIG. 2B is a detailed diagram illustrating the system of FIG. 2B, with the same switches in the selection position of a resistor; FIG. 3 is a detailed diagram illustrating a space propulsion system according to a second embodiment; FIG. 4 is a detailed diagram illustrating a space propulsion system according to a third embodiment; and FIG. 5 is a detailed diagram illustrating a space propulsion system according to a fourth embodiment.
    [0009] DETAILED DESCRIPTION OF THE INVENTION FIG. 1 illustrates a spacecraft 10, more specifically a satellite, equipped with an orientation and trajectory control system serving to maintain the orbit and the orientation of the spacecraft relative to the body. that he orbits. For this, the steering and trajectory control system comprises, apart from at least one sensor 11 intended to determine the real orientation and trajectory of the spacecraft, and a control unit 12, connected to the sensor 11 and intended to determine the desired orientation and trajectory and the maneuvers to be performed to achieve the desired orientation and trajectory from the actual orientation and trajectory determined by the at least one sensor 11, operating means connected to the control unit 12, and capable of exerting forces and torques on the spacecraft 10 to perform said maneuvers. In the illustrated example, these operating means comprise in particular a space propulsion system 100, although other operating means, such as inertial devices such as reaction wheels, or devices using the pressure of solar radiation, can be considered in addition to this space propulsion system 100. Moreover, the spacecraft 10 further comprises a power source 13, in the form of photovoltaic panels in the example illustrated, although other power sources electrical, for example batteries, fuel cells or thermoelectric generators, are also possible in addition or alternatively to these photovoltaic panels. This power source 13 is connected to the various electrical consumers of the spacecraft through a main power bus 14. In addition, the spacecraft 10 also comprises at least one propellant reservoir 15, such as by example of Xenon. Figures 2A and 2B illustrate the spatial propulsion system 100 according to a first embodiment. This space propulsion system 100 comprises an electrostatic thruster 101 and a resistor 102. In addition, it also comprises a power supply circuit 103, and a propellant fluid supply circuit 104, both connected to the two thrusters for their supply, respectively, electricity and propellant. The power supply circuit 103 is connected to the power supply source 13 of the spacecraft 10 through the bus 14. The propellant fluid supply circuit 104 is connected to the tank 15. The electrostatic propellant 101, which is more specifically a Hall effect thruster, comprises a channel 150 with an annular section, closed at its upstream end and open at its downstream end, an anode 151 located at the upstream end of the channel 150, a transmitting cathode 152, located downstream the downstream end of the channel 150 and equipped with at least one heating element 153, the electromagnets 154, located radially inside and outside the channel 150, and the injectors 155 of propellant fluid, located at the upstream end of the channel 150.
    [0010] The resistor 102 is simpler, mainly comprising at least one propellant fluid injector 160, a heating element 161, and a nozzle 162. As can also be seen in FIGS. 2A and 2B, the propellant fluid supply circuit 104 comprises a line 105 for supplying the electrostatic propellant 101 with propellant fluid connected to the injectors 155 of the electrostatic thruster 101, and a line 106 for supplying the resistor 102 with propellant fluid connected to the injector 162 of the resistor 102. On the line 105 An electrostatic propellant booster 101 is supplied with propellant gas regulator 107, while a booster pressure regulator 102 of the resistor 102 is installed on the line 106. These pressure regulators 107, 108 thus make it possible to ensure substantially constant supply pressures of the two thrusters, even when the pressure of the reservoir 15 varies greatly. Although in the embodiment illustrated two different pressure regulators are used to obtain different supply pressures, it would also be possible to use a single common pressure regulator to provide the same pressure to the two thrusters.
    [0011] A flow regulator 109 is also installed on the line 105 for supplying the electrostatic propellant 101 with propellant gas, downstream of the pressure regulator 107 but also upstream of the injectors 155 of propellant fluid in the electrostatic thruster 101. This flow regulator 109 includes an on-off valve 110 and a thermostrictor 111 installed in series to respectively control the supply of propellant electrostatic propellant 101 and regulate its flow. Moreover, the propellant fluid supply circuit 104 also comprises a bypass 171 connecting the line 105 downstream of the flow regulator 109 to the cathode 152, in order to bring a very limited flow rate of gas to this cathode 152. which is a hollow cathode, so as to facilitate the emission of electrons from the cathode 152, as well as its cooling. A throttling 172 on this bypass 171 restricts the flow of propellant gas supplied to the cathode relative to that injected through the injectors 155.
    [0012] The propellant fluid supply circuit 104 also comprises a valve 112 for supplying the propellant gas resistor 102, directly integrated into the resistor 102, upstream of the injector 160, in the illustrated embodiment, even though it is also alternatively possible to install it on the line 106, between the pressure regulator 108 and the resistor 102.
    [0013] The power supply circuit 103 comprises an electrical conditioning unit 113 (or PPU, according to the acronym of "Power Processing Unit") with a thruster selection unit 114 (or TSU, according to the acronym "Thruster" Selection Unit "). Although in the illustrated embodiment the selection unit 114 is integrated in the conditioning unit 113, it is also conceivable to arrange it outside thereof. In this case, such an external unit for selecting thrusters can receive the acronym ETSU ("External Thruster Selection Unit"). The electrical conditioning unit 113 also comprises a limiter 115, inverters 116, a control interface 117, a sequencer 118, and a DC voltage converter 119. Furthermore, the electrical conditioning unit 113 comprises in addition to a heater power supply current regulator 120, a voltage regulator 121 of VD + ND- and current supply ID of the anode 151 and the cathode 152, a regulator 122 of the feed current Im electromagnet, electric ignition pulse regulators 123, a valve control regulator 124 and a thermoregulator control current regulator 125. For their power supply, these regulators 120 to 125 are all connected to a first power input 126 of the conditioning unit 113 through the inverters 116. The control interface 117 and the sequencer 118 are connected to a second one. supply inlet 127 of the conditioning unit 113 through the converter 119 for their own power supply and, through a control input 128, to the control unit 12 of the steering control system and trajectory. They are also connected to regulators 120 to 125 to control their operation. The selection unit 114 comprises a set of switches each connected to one of the outputs of the regulators 120 to 125 through a corresponding power supply or control line. Thus, the regulator 120 is connected to the switch 114-1 by a first power supply line 131, the regulator 121 to the dual switch 114-2 by a second and a third power supply line 132+, 132-, the regulator 122 to the switch 114-3 by a fourth power supply line 133, the regulator 123 to the switch 114-4 by a fifth power supply line 134, the regulator 124 to the switch 114-5 by a control line 135 valve opening, and that of the regulator 125 to the switch 114-6 by a line 136 of the thermorestrictor control. Each switch can switch between at least a first contact A and at least a second contact B, and the selection unit 114 is connected to the control unit 12 so as to allow it to control the simultaneous switching of each switch. In the illustrated embodiment, each switch contact 114-1 to 114-4 of a first group is connected to an electrical consumer of the electrostatic thruster 101. Thus, the switch contact 114-1 is connected to the switch. heating element 153 of the emitting cathode 152, and those of the switches 114-3 to 114-4 to, respectively, the electromagnets 154 and ignition means (not shown) of the electrostatic thruster 101. In the illustrated embodiment, each of these electrical consumers is grounded, so that a simple switch and a single line of power supply is enough to ensure the supply of each of them. However, it is also possible to isolate each of these electrical consumers and to do without grounding by using return lines and double switches connected not only to the outgoing lines, but also to the return lines to switch or interrupt their contact. . Thus, in the illustrated embodiment, one of the contacts A of the dual switch 114-2 is connected to the cathode 152 through a filter device 170 and can be connected by the switch 114-2 to the power supply line 132. , with negative polarity, and the other of the contacts A of the double switch 114-2 is connected to the anode 151 through the same filtering device 170 and can be connected by the switch 114-2 to the supply line 132 +, positive polarity. In addition, each contact A switches 114-5 and 114-6 of a second group is connected to the flow controller 109 of the line 105 for supplying the thermostatic thruster 101 with propellant. In particular, the contact A of the switch 114-5 is connected to the valve 110, while that of the switch 114-6 is connected to the thermoregulator 111. On the other hand, in the embodiment illustrated, the contact B of the switch 114 -1 and that of the switch 114-5 are respectively connected to the heating element 161 and the valve 112 of the resistor 102. Thus, in operation, the electrical conditioning unit 113 can provide the power supply and control the power supply. supply of propellant fluid is the electrostatic propellant 101 or the resistor 102, according to a selection operated through the unit 114 for selecting propellants. In this way, when the switches 114-1 to 114-6 connect the power supply lines 131, 132-F, 132-, 133 and 134 to the electrostatic thruster 101 and the control lines 135 and 136 to the flow controller 109 as shown in FIG. 2A, the electrostatic thruster 101 can be activated and controlled by the control unit 12 of the spacecraft 10, through the electrical conditioning unit 113. In particular, signals from the control unit 12 are transmitted to the regulators 120 to 125 through the control interface 117 and the sequencer 118, and in this case control the power supply of the various electrical consumers of the electrostatic thruster 101 through the regulators 120 to 123 of on the one hand, and on the other hand, through the regulators 124 and 125 and the flow regulator 109, supplying the electrostatic propellant 101 with propellant fluid.
    [0014] On the other hand, when the switches 114-1 to 114-6 switch to the contacts B, as illustrated in FIG. 2B, the first power supply line 131 is connected to the heating element 161 of the resistor 102, while the line 135 of the valve opening control is connected to the valve 112 of the resistor 102. In this manner, signals from the control unit 12 and transmitted to the controllers 120 and 124 through the control interface 117 and the sequencer 118 can then control the electrical supply of the heating element 161 of the resistor 102 through the regulator 120 on the one hand, and on the other hand, through the regulator 124 and the valve 112, the supply of the resistor 102 in propulsive fluid.
    [0015] The space propulsion system 100 according to this first embodiment can therefore operate in a first propulsion mode, with a high specific impulse but a low thrust, by selecting the electrostatic thruster 101 through the selection unit 114, or in a second mode. propulsion system, with a lower specific impulse, by selecting the resistor 102 through the selection unit 114. Although in this first embodiment the supply of the electrostatic propellant 101 fluid is carried through a pressure regulator and a flow regulator comprising a valve and a thermostrictor, in other embodiments the supply of the electrostatic propellant in fluid can be carried out through a joint pressure and flow control unit comprising two damsels arranged in series. Thanks to the impedance of the propellant fluid supply circuit, in particular between the two on-off valves, it is possible to regulate the pressure and flow rate of the propulsive fluid supplied to the electrostatic propulsion unit by controlling pulsations of the two all-or-nothing valves. -nothing. The pressure of the propulsive fluid supplied to the resistor can also be controlled in the same way. Thus, in a second embodiment illustrated in FIG. 3, the pressure and flow regulators on the first fluid gas supply line of the propulsion system according to the first embodiment can be replaced by a single regulator 109 'of pressure and flow comprising two on-off valves 110 ', 111' arranged in series on the line 105 for supplying the electrostatic propellant 101 with propellant gas. In this second embodiment, the valve of the resistojet and the corresponding pressure regulator are also replaced by a single regulator 112 'pressure and flow also comprising two all-or-nothing valves 112'a, 112'b arranged in series on the feed line 106 of the resistor 102 in propellant fluid. In the power supply conditioning unit 113, the thermoregulator control current regulator LU of the first embodiment is replaced by a second valve opening control regulator 125 '. The other elements of the system according to this second embodiment are similar to those of the first embodiment and accordingly receive the same reference numbers in FIG. 3 as in FIGS. 2A and 2B.
    [0016] Thus, during the operation of the space propulsion system 100 according to this second embodiment, when the electrostatic thruster 101 is selected through the thruster selection unit 114 and its switches 114-1 to 114-6, signals from of the control unit 12 and transmitted to the regulators 124 and 125 through the control interface 117 and the sequencer 118 control the valves 110'411 'of the regulator 109' in order to regulate the supply of the electrostatic propellant 101 with propellant . On the other hand, when the resistor 102 is selected through the thruster selection unit 114 and its switches 114-1 to 114-6, these same signals may control the 112 'and 112' b valves of the regulator 112. in order to regulate the supply of the resistor 102 to propellant fluid. For the remainder, the operation of the space propulsion system 100 according to this second embodiment is similar to that of the first embodiment, in particular as regards the regulation of the electric power supply of the electrostatic thruster 101 and the resistor 102, and selecting two different modes of propulsion. Although, in the two previous embodiments, the electrical supply of the heating elements of the resistor and the cathode of the electrostatic propellant, respectively, passes through the electrical conditioning unit, and in particular through one of the Inverters, it is also possible to bypass the electrical conditioning unit for the power supply of these elements. Indeed, the operating voltage of the heating elements of the two thrusters may be close to, and even equal to, that of the main supply bus, thereby allowing their power supply directly from it. Thus, in a third embodiment illustrated in FIG. 4, the first power supply line 131 comes from a switch 120 "directly connected to the main supply bus 14 and to the control unit 12 of the spacecraft 10. Although, in the illustrated embodiment, the switch 120 "is separate and distinct from the electrical conditioning unit 113, it is also conceivable to integrate it therein. Moreover, in the embodiment illustrated, the thruster selection unit 114 is also external to the electrical conditioning unit 113, although their integration can also be envisaged. The other elements of the system according to this third embodiment are however similar to those of the first embodiment and accordingly receive the same reference numerals in Figure 4 as in Figures 2A and 2B. In this way, during the operation of the space propulsion system 100 according to this third embodiment, when the electrostatic thruster 101 is selected through the thruster selection unit 114 and its switches 114-1 to 114-6, signals transmitted by the control unit 12 to the switch 120 "can control pulsations of current on the first power supply line 131 to regulate the operation of the heating element 153 of the emitting cathode 152 of the electrostatic thruster 101. D On the other hand, when the resistor 102 is selected through the thruster selection unit 114 and its switches 114-1 to 114-6, these same pulses may regulate the operation of the heater element 161 of the resistor 102. For the remaining, the operation of the system 100 of space propulsion according to this third embodiment is similar to that of the first embodiment, in particular as regards the regulation of the power supply of the electrostatic thruster 101 and the resistor 102 in propellant fluid, and the selection of two different modes of propulsion. Although, in the three previous embodiments, the space propulsion system comprises only a single electrostatic propellant and a single resistor, the same principles are equally applicable to systems comprising pluralities of electrostatic thrusters and resistors. Thus, in a fourth embodiment illustrated in FIG. 5, the space propulsion system 100 comprises two electrostatic thrusters 101 and two resistors 102, arranged for example in two pairs of thrusters, each formed by an electrostatic thruster 101 and a resistor 102. , the thrusters of one of these pairs being oriented in opposite directions to those of the other. The two electrostatic thrusters 101 are connected to a single regulator 107 of supply pressure of the electrostatic propellants 101 in propellant gas by lines 105 for supplying propellant fluid, while the two resistors 102 are also connected to a single regulator 108 supply pressure of the resistors 102 in propellant gas by other lines 106 for supplying propellant. On the other hand, individual flow regulators 109 are installed on each propulsive fluid supply line 105 of the electrostatic thrusters 101 in order to regulate separately the supply flow rate of each electrostatic propellant 101 in propellant fluid. The propellant fluid supply circuit 104 also comprises a propellant gas supply valve 112 for each resistor 102. In addition, this space propulsion system 100 also comprises two external units 114 ', 114 "of propellant selection in addition. a thruster selection unit 114 integrated in the electrical conditioning unit 113. The three thruster selection units 114, 114, 114 "are connected to the control unit 12 of the spacecraft 10 to control their respective switches. 114-1 to 114-6, 114'-1 to 114'-6, and 114 "-1 to 114" -6. The contacts A of the thruster selection unit 114 are connected to the electrostatic thruster 101 or to the resistor 102 of a first of said pairs of thrusters through the first external selection unit 114 ', while the contacts B of the unit The other elements of the system according to this fourth embodiment are connected to the electrostatic thruster 101 or to the resistor 102 of the second of said pairs of thrusters through the second external selection unit 114 ". embodiment and accordingly receive the same reference numerals in Fig. 5 as in Figs. 2A and 2B Thus, in operation, the electrical conditioning unit 113 can provide power and control the supply of propellant fluid. either a thruster of the first pair or a thruster of the second pair, following a selection op created through the thruster selection unit 114. If the first pair of thrusters is selected through the selection unit 114, the selection between the electrostatic thruster 101 and the resistor 102 of this first pair can be made through the first external selection unit 114 'in a manner analogous to the selection of thrusters in the previous embodiments. Similarly, if the second pair of thrusters is selected through the selection unit 114, the selection between the electrostatic thruster 101 and the resistor 102 of this second pair can be made through the second external selection unit 114. similarly to the selection of thrusters in the previous embodiments, thus, through the switches of the three selection units 114, 114 'and 114 ", it is possible to choose between two directions of propulsion, and between two modes of propulsion. in each direction. For the remainder, the operation of the system 100 of space propulsion according to this fourth embodiment is similar to that of the first embodiment, in particular with regard to the regulation of the propulsion power propulsive propulsion and electricity. Although the present invention has been described with reference to a specific exemplary embodiment, 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. In particular, the eigen characteristics of the second and / or third embodiments could also be adapted to a system comprising a plurality of propeller and thruster selection units of each type, as in the fourth embodiment. Furthermore, although the system according to the fourth embodiment comprises only two pairs of different types of thrusters, it would also be possible to incorporate a greater number of pairs therein. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.
    权利要求:
    Claims (13)
    [0001]
    REVENDICATIONS1. A space propulsion system (100) comprising at least: an electrostatic propellant (101) with at least a first electrical consumer; a resist (102); a propellant fluid supply circuit (104); and a power supply circuit (103) having at least a first power supply line (131) and a first switch (114-1, 114'-1,114 "-1) for selecting between connecting said first line (131) supplying power to the resistor (102) and connecting said first power supply line (131) to said first electrical consumer of the electrostatic thruster (101).
    [0002]
    The space propulsion system (100) of claim 1, wherein said first electrical consumer comprises a heater (153) of an emitter cathode (152) of said electrostatic booster (101).
    [0003]
    Spatial space propulsion system (100) according to any one of the preceding claims, wherein said propellant fluid supply circuit (104) comprises at least one electrostatic propellant supply valve (110, 110 ') (101). ) and at least one valve (112) for supplying the resistor (102).
    [0004]
    The space propulsion system (100) according to claim 3, further comprising at least one valve opening control line (135) and a second switch (114-5,114'-5,114-5 ") for selecting between connecting said valve opening control line (135) to the electrostatic thruster (101) valve (110, 110 ') and connecting said valve opening control line (135) to the at least one supply valve (112) of the resistor (102).
    [0005]
    A space propulsion system (100) according to any one of the preceding claims, wherein said power supply circuit (103) comprises at least one propellant selection unit (114, 114, 114 ") in which said at least one first switch (114-1,114'-1,114 "-1) is integrated. 35
    [0006]
    The space propulsion system (100) according to any one of the preceding claims, wherein said power supply circuit (103) further comprises at least one electrical conditioning unit (113) adapted to supply at least one other consumer electrostatic thruster (101) at a substantially higher voltage than the first electrical consumer.
    [0007]
    The space propulsion system (100) according to any one of the preceding claims, wherein said electrostatic propellant (101) is a Hall effect booster.
    [0008]
    A space propulsion system (100) according to any one of the preceding claims, comprising a plurality of electrostatic thrusters (101).
    [0009]
    The space propulsion system (100) of claim 9, wherein said propellant supply circuit (104) comprises at least one pressure regulator (107) common to a plurality of said electrostatic propellants (101).
    [0010]
    The space propulsion system (100) according to claim 9 or 10, wherein said propellant supply circuit (104) comprises a pressure regulator (107) individually for at least one of said electrostatic thrusters (101).
    [0011]
    An orientation and / or trajectory control system comprising a space propulsion system (100) as claimed in any one of the preceding claims.
    [0012]
    Spacecraft (10) comprising a space propulsion system according to any one of claims 1 to 11.
    [0013]
    13. A space propulsion method comprising a step of switching between an electrostatic thruster (101) and a resistor (102), wherein a first switch (114-1, 114'-1,114 "-1) is used to connect a first line ( 131) to the resistor (102) or a first electrical consumer of the electrostatic thruster (101), for selecting a first propulsion mode, wherein the resistor (102) is activated, or a second mode of propulsion, wherein the electrostatic propellant (101) is activated.
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    法律状态:
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    2016-02-05| PLSC| Publication of the preliminary search report|Effective date: 20160205 |
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    2017-05-02| PLFP| Fee payment|Year of fee payment: 4 |
    2017-11-10| CD| Change of name or company name|Owner name: SNECMA, FR Effective date: 20170713 |
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1457371A|FR3024436B1|2014-07-30|2014-07-30|SYSTEM AND METHOD FOR SPACE PROPULSION|
    FR1457371|2014-07-30|FR1457371A| FR3024436B1|2014-07-30|2014-07-30|SYSTEM AND METHOD FOR SPACE PROPULSION|
    US15/329,470| US20170210493A1|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    JP2017504808A| JP6672260B2|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    PCT/FR2015/052067| WO2016016563A1|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    RU2017106191A| RU2684968C2|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    CN201580041827.XA| CN106574607B|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    EP15756968.2A| EP3174795B1|2014-07-30|2015-07-27|Spacecraft propulsion system and method|
    IL250167A| IL250167A|2014-07-30|2017-01-17|Spacecraft propulsion system and method|
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