Engine for a space launch vehicle, space launch vehicle and method for implementing such an engine

Abstract

The invention relates to an engine for a space launch vehicle with at least one stage extending along a longitudinal axis (Z), the engine comprising: - a power group configured to pressurize at least one propellant to a predetermined pressure; - 2n combustion chambers (A, A', B, B') having a fixed spatial orientation, where n≥2, arranged in diametrically opposed and fluidically paired pairs; - a fluid circuit for distributing and controlling the at least one propellant, the fluid circuit being configured to distribute the at least one propellant to the chambers (A, A', B, B') by controlling the propellant flow distribution between the paired chambers in order to inject, for a pair of chambers, a flow rate Q+α into a first chamber (A) and a flow rate Q-α into a second chamber (A'), where Q is the average flow rate of the injected propellant (E1) and α is a value of variation of this flow rate.

Description

Description Title of the invention: Engine for a space launcher, space launcher and method for implementing such an engine Technical Field

[0001] This presentation concerns a space launcher engine, a space launcher, and a method for implementing such a space launcher engine. Previous technique

[0002] In the space sector, rocket or space launcher engines generally include a power unit that supplies propellants to one or more combustion chambers in which thrust is generated.

[0003] As is known, the overall thrust generated by the combustion chamber(s) must be able to be spatially orientable, in pitch and yaw, in order to direct the launcher along the desired trajectory.

[0004] We know of space launch vehicle stage architectures that include one or more engines equipped with actuators to spatially orient the combustion chambers and thus the overall thrust they generate. However, actuator activation requires these heavy components to be carried onboard, also necessitates auxiliary components to move them (electronic, hydraulic, pneumatic systems, etc.), and results in significant size and weight, which can be detrimental to a launch vehicle. Description of the invention

[0005] In view of the above, there is clearly a need to reduce the size and / or weight of a powered stage of a space launcher.

[0006] One embodiment of the invention relates to a space launcher engine with at least one stage extending along a longitudinal axis Z, said engine including: -a power unit configured to pressurize at least one propellant to a predetermined pressure, the power unit comprising a pressurization unit for said at least one propellant which is driven by a pre-combustion chamber or by a gas generator or by the outlet of a heat exchanger, -2n combustion chambers of fixed spatial orientation, with n>2, which are arranged, according to a view taken in a plane perpendicular to the longitudinal axis, two by two diametrically opposite each other on a circumference of a circle centered on the longitudinal axis Z and fluidly matched, each of the combustion chambers being capable of generating on command a thrust directed along the longitudinal axis Z, -a fluidic distribution and control circuit for said at least one pressurized propellant which is configured to distribute said at least one pressurized propellant into the 2n combustion chambers by controlling the flow distribution of said at least one pressurized propellant between the fluidly matched combustion chambers so as to inject, for at least one pair of fluidly matched combustion chambers, a flow Q+o of said at least one propellant E1 into a first combustion chamber of said at least one pair of fluidly matched combustion chambers and a flow Q-α of said at least one propellant E1 into a second combustion chamber of said at least one pair of fluidly matched combustion chambers, where Q is the average flow of said at least one pressurized propellant E1 from the power group and α a value of variation of this flow.

[0007] The launcher engine configuration defined above allows for the generation of a total thrust, with combustion chambers of fixed spatial orientation. The overall thrust generated by these chambers can vary in the X, Y space (perpendicular to the longitudinal Z axis) depending on the distribution of the propellant flow rate supplying the paired, fluidly coupled chambers. This configuration eliminates the need for actuators to change the chamber orientation, thereby reducing the overall size and weight of the stage. Onboard. Note that the average flow rate Q corresponds to the flow rate for a "straight" thrust of the engine, that is, a thrust generated solely along the longitudinal axis Z. A pre-combustion chamber, a gas generator, or an outlet of a heat exchanger each constitutes a power source (e.g., thermodynamic) for the power unit's pressurization system. This source is supplied with propellant, which is, for example, heated by cooling the gases from the engine's combustion chambers.

[0008] Depending on other possible characteristics: -the power unit and the fluidic distribution and control circuit are configured respectively to pressurize a single propellant E1 and to distribute it into each of the 2n combustion chambers; - The fluid circuit for distributing and controlling the pressurized propellant is configured to distribute the pressurized propellant E1 into the 2n combustion chambers by controlling the flow distribution of the pressurized propellant between the 2n fluidly matched combustion chambers so as to inject, for at least two pairs of combustion chambers: - a flow rate Q1A+o of propellant E1 in a first combustion chamber of a first pair and a flow rate Q1A-o of propellant E1 in a second combustion chamber of the first pair, where Q1A is the flow rate of propellant E1 and has a value of variation with respect to this propellant flow rate, - a flow rate Q1 B+y of propellant E1 in a third combustion chamber of a second pair and a flow rate Q1B-y of propellant E1 in a fourth combustion chamber of the second pair, where Q1 B is the flow rate of propellant E1 and y is a value of variation with respect to this propellant flow rate; -the fluid circuit for the distribution and control of the pressurized propellant E1 includes fluid lines equipped with distribution and control valves which are configured to connect the power unit to the 2n combustion chambers; - the fluid circuit for the distribution and control of the pressurized propellant (E1) includes a distribution and control valve (VE1a, VE1b) for the propellant flow rate for each of the pairs of fluidly matched combustion chambers; -the power unit and the fluidic distribution and control circuit are configured respectively to pressurize two propellants E1, E2 and to distribute them into each of the 2n combustion chambers; -The fluidic circuit for the distribution and control of pressurized propellants is configured to distribute the pressurized propellants into the 2n combustion chambers by controlling the distribution of pressurized propellant flow rates between the fluidically paired combustion chambers so as to inject, for at least two pairs of combustion chambers: a flow rate Q1A+a of a first propellant E1 in a first combustion chamber A of a first pair and a flow rate of the first propellant Q1 Aa in a second combustion chamber A' of the first pair, a flow rate Q2A+P of a second propellant E2 in the first combustion chamber A of the first pair and a flow rate Q2A-P of the second propellant E2 in the second combustion chamber A' of the first pair, where Q1A and Q2A are respectively the flow rates of the first propellant E1 and the second propellant E2 and a, p are the values ​​of variation relative to these flow rates of the first and second propellants, a flow rate Q1 B+y of the first propellant E1 in a third combustion chamber B of a second pair and a flow rate Q1 By of the first propellant in a fourth combustion chamber B' of the second pair, a flow rate Q1 By of the second propellant E2 in the third combustion chamber B of the second pair and a flow rate Q2B-5 of the second propellant E2 in the fourth combustion chamber B' of the second pair, where Q1 B and Q2B are respectively the flow rates of the first propellant E1 and the second propellant E2 and y, 5 are the values ​​of variation with respect to these flow rates of the first and second propellants; -the fluidic circuit for the distribution and control of pressurized propellants includes fluid lines equipped with distribution and control valves which are configured to connect the power unit to the 2n combustion chambers; -The fluidic circuit for the distribution and control of pressurized propellants includes, for each of the two propellants, a valve for distributing and controlling the propellant flow rate in each of the two combustion chambers of the same pair of matched combustion chambers; -the distribution and control valves are electrically or hydraulically operated; -The distribution and control valves are of the linear or rotary type; -The distribution and control valves are three-way valves; -the power group includes a distribution and control valve for each pressurized propellant intended to supply each pair of fluidically matched combustion chambers; this simplifies the configuration of the fluidic circuit for the distribution and control of said at least one pressurized propellant; -the power unit includes at least one VLO control valve allowing control of a propellant supply flow to the power unit; this allows the drive power of the pressurization unit (for example a turbopump) to be modified on command and therefore the overall thrust of the engine to be modified by acting on the 2n chambers simultaneously.

[0009] Another embodiment relates to a space launcher with at least one stage extending along a longitudinal axis Z, the launcher comprising: -a space launcher engine as briefly described above, -at least one control unit which is configured to control the engine's distribution and control circuit.

[0010] Depending on other possible characteristics: -said at least one control unit is configured to control the distribution and control valves of the engine's propellant distribution and flow control circuit; -said at least one control unit is configured to control the distribution and flow control valves of said at least one propellant according to an overall thrust setpoint; -said at least one control unit is configured to control the propellant distribution and flow control valves according to an overall thrust setpoint and a mixture ratio between propellants for matched combustion chambers.

[0011] Another embodiment relates to a method of implementing an engine for a space launcher with at least one stage extending along a longitudinal axis Z, said engine comprising: -2n combustion chambers of fixed spatial orientation, with n>2, which are arranged, according to a view taken in a plane perpendicular to the longitudinal axis Z, two by two diametrically opposite each other on a circumference of a circle centered on the longitudinal axis Z and fluidly matched, each of the combustion chambers being capable of generating on command a thrust directed along the longitudinal axis Z, -a fluidic circuit for the distribution and control of at least one propellant in the 2n combustion chambers, the process comprising: -the pressurization of said at least one propellant E1 to a predetermined pressure from a controlled power unit, -the control of the flow distribution of said at least one pressurized propellant between the fluidly matched combustion chambers so as to inject, for at least one pair of fluidly matched combustion chambers, a flow Q+a of said at least one propellant E1 into a first combustion chamber of said at least one pair of fluidly matched combustion chambers and a flow Qa of said at least one propellant E1 into a second combustion chamber of said at least one pair of fluidly matched combustion chambers, where Q is the average flow of said at least one pressurized propellant E1 and a is a value of variation of this flow.

[0012] The power group can be controlled by controlling a propellant supply rate from a power source.

[0013] Since the process has the same advantages and characteristics as the launcher engine briefly described above, they will not be repeated here. Brief description of the drawings

[0014] The purpose and advantages of this presentation will be better understood upon reading the detailed description below of different methods of The implementations given are non-exhaustive examples. This description refers to the attached figure pages, on which:

[0015] [Fig. 1] Figure 1 represents a possible embodiment of a space launcher stage engine.

[0016] [Fig. 2] Figure 2 represents one possible configuration of the four combustion chambers of Figure 1.

[0017] [Fig. 3] Figure 3 is a schematic representation of a control unit configured to cause a change in the thrust direction of the engine of Figure 1.

[0018] [Fig. 4A] Figure 4A illustrates in more detail an aspect of the operation of the MVA multi-variable controller of Figure 3.

[0019] [Fig. 4B] Figure 4B illustrates in detail and enlarged the VE2a valve of figure 3.

[0020] [Fig. 5] Figure 5 represents another possible embodiment of a space launcher stage engine. Description of the implementation methods

[0021] Figure 1 schematically and partially represents a pump engine 10 of a stage of a one or more stage space launcher. The stage considered here is, for example, the second stage of the launcher.

[0022] The stage portion shown in Figure 1 comprises a pump engine 10 including a power unit configured to pressurize two propellants, a fuel C1 (known as 'fuel' in Anglo-Saxon terminology) and an oxidizer C2 (known as 'oxidant' in Anglo-Saxon terminology), to a predetermined pressure corresponding to a specific thrust operating point of the launch vehicle. The following description pertains to a staged combustion engine shown in Figure 1 but also applies to other engine types in which the power unit can be configured for a gas generator engine or an expansion engine. Depending on the type of power unit, the oxidizer C2 could be, for example, liquid oxygen (LOX), and the fuel C1 could be either methane (CH4) or hydrogen. Liquid LH2 is a fuel-oxidizer mixture with a higher proportion of fuel than oxidizer, or a higher proportion of oxidizer than fuel. The supply flows of unpressurized propellants C1 and C2 from tanks (not shown) are identified in the upper part of the diagram. In the remainder of this presentation, we will discuss, in general terms, pressurized propellants exiting the power unit, propellant E1 and propellant E2.

[0023] Engine 10 also includes several combustion chambers, each with a fixed spatial orientation.

[0024] More specifically, the number of chambers is even and equal to 2n, with n > 2. In this embodiment, the chambers are fluidly paired. In the example embodiment illustrated in Figure 1, there are four chambers, named A, A', B, and B'. Chambers A and A' are fluidly paired with each other, while chambers B and B' are fluidly paired with each other.

[0025] Figure 2 shows the geometric arrangement in space [X, Y] (plane perpendicular to the longitudinal axis Z along which the launcher extends) in which these four chambers are arranged in the considered stage of the space launcher. This arrangement is not shown in the diagram in Figure 1, which is a schematic diagram.

[0026] As shown in Figure 2, the combustion chambers A, A', B, B' are arranged in the X, Y plane, here on a circumference of a circle C, although this arrangement may vary. The combustion chambers are arranged in pairs diametrically opposite each other on the circumference of circle C: A and A' on one side, and B and B' on the other.

[0027] The power group in Figure 1 includes in particular a unit 12 for pressurizing the two propellants C1 and C2, which will be described in more detail below.

[0028] In Figure 1, the engine 10 also includes a fluidic circuit for the distribution and control of pressurized propellants, labeled E1 and E2 in Figure 1, which is configured to distribute the two types of pressurized propellants. in each of the 2n combustion chambers (here A, A', B and B') by controlling the distribution of pressurized propellant flow rates between the fluidically matched combustion chambers, namely A, A', on the one hand, and B, B', on the other. This circuit generally comprises: - a plurality of pipes or conduits distributing, on the one hand, the pressurized propellant E1 from the pressurization unit 12 to the combustion chambers A, A', B, B' and, on the other hand, the pressurized propellant E2 from the pressurization unit 12 to the combustion chambers A, A', B, B'; - several distribution and control valves VE1a, VE1b, VE2a, VE2b which are connected / linked to the plurality of distribution lines and which each allow the distribution and control of pressurized propellant flows to the paired combustion chambers.

[0029] The pressure-generating unit 12 of the power group comprises, for example, a turbine 14, referred to as the low-pressure pump, and two high-pressure pumps 16 and 18 mounted one behind the other. The turbine 14 is connected via an output shaft 14a to an inlet of the pump 16, whose output shaft 16a is connected to the inlet of the pump 18. The pump 16 is supplied at its inlet by the propellant flow C1 from a reservoir not shown in the figure. The pump 16 is connected at its outlet to a cooling zone located downstream of the combustion chambers A, A', B, B', via a common connection LO, equipped with a control valve VLO (controllable), which carries a pressurized propellant flow Cp1 (at a pressure higher than that of the pressurized propellant flow E1), from which respective cooling fluid lines or connections L1, L2, L3, L4 branch off.Pump 18 is supplied with propellant C2 from a reservoir not shown in the figure and is connected at its outlet, via distribution and control valves VE2a and VE2b, to the inlets of combustion chambers A, A', B, and B', to which the pressurized propellant E2 is delivered. Turbine 14 is connected at its outlet, via distribution and control valves VE1a and VE1b, to the inlets of combustion chambers A, A', B, and B', to which the pressurized propellant E1 is delivered.

[0030] The power unit also includes, as a power source in this embodiment, a pre-combustion chamber 20 (for a staged combustion engine as is the case here) which receives propellant flows from the pressurized and heated oxidizer from the cooling lines of the cooling zone of each of the combustion chambers via respective lines or connections l1, l2, l3, l4 and associated valves (not shown), or, for example, from a control valve such as the aforementioned VL0 valve. The VL0 valve allows control / control of the propellant supply flow to the pre-combustion chamber 20 and therefore to the pressurization unit 12. This makes it possible to modify the drive power of the pressurization unit 12 (here a turbopump) and thus modulate the total thrust of the engine via the thrusts of each of its combustion chambers.It should be noted that for a gas generator engine, the power source would be a gas generator, not a pre-combustion chamber. In other configurations, the outlet of a heat exchanger could provide the necessary energy (thermodynamic energy) for propellant pressurization. For an expansion engine, there is neither a pre-combustion chamber nor a gas generator; the connections l1, l2, l3, and l4 directly supply the turbine 14. However, the VL0 valve can control the turbine 14 via the distribution lines that pass through the combustion chamber cooling circuits.

[0031] The pre-chamber 20 also receives the pressurized propellant E2 (oxidizer) from pump 18, supplied by valve V5. The pre-chamber 20 is connected to the turbine 14 via a fluid line L5. The function of the pre-chamber 20 is to generate an enthalpy flow, in this case hot gases from the combustion of the pressurized propellant Cp1 from pump 16 and the pressurized propellant E2. This generated flow is transmitted to the turbine 14 via the line L5 to drive the turbine 14 in rotation and, consequently, to drive the pumps 16 and 18 in rotation to pressurize the propellants C1 and C2. This pre-chamber 20, together with the pressurization unit 12, together form the engine's power unit.

[0032] The circuit includes, at the outlet of the turbine 14, a common fluidic conduit or link 14.0 that distributes the pressurized propellant flow E1 into two fluidic conduits or links 14.1 and 14.2, which carry fractions of this pressurized propellant flow E1 respectively to the two valves VE1a and VE1b for distributing and controlling the pressurized propellant E1. These valves then distribute the relevant fraction of the pressurized propellant flow E1 to the four combustion chambers A, A', B, and B'. Thus, valve VE1a distributes propellant E1 in a controlled manner to chambers A and A', and valve VE1b distributes propellant E1 in a controlled manner to chambers B and B'.

[0033] Furthermore, the circuit also includes, at the outlet of pump 18, a common fluid line or connection 18.0 distributing the pressurized propellant flow E2 into fluid lines or connections 18.1 and 18.2, which carry fractions of this pressurized propellant flow E2 respectively to the two valves VE2a and VE2b for distributing and controlling the pressurized propellant E2. These valves then distribute the relevant fraction of the pressurized propellant flow E2 to the four combustion chambers A, A', B, and B'. Thus, valve VE2a distributes propellant E2 in a controlled manner to chambers A and A', and valve VE2b distributes propellant E2 in a controlled manner to chambers B and B'.

[0034] The valves VE1 a, VE1 b, VE2a, VE2b are each appropriately electronically controlled by a control unit described later, which distributes the flow rates or fluxes of pressurized propellants between the fluidly matched chambers (A and A', on the one hand, and B and B' on the other) in order to generate thrust with the desired spatial orientation.

[0035] In this example embodiment, the valves VE1a, VE1b, VE2a, and VE2b are preferably three-way valves. However, each of these valves can alternatively be replaced by a set of simple two-way valves performing the same function.

[0036] The valves here are electrically controlled but can alternatively be hydraulically controlled. The valves can be of the linear or rotary type.

[0037] It should be noted that a set of simple shut-off valves V1 to V5 is provided on the pipes to allow isolation, if necessary, of certain parts of the circuit and to allow management of start-up and stop-down transients.

[0038] According to the embodiment shown in Figure 1 (staged combustion), the propellant flow C1 is supplied to pump 16, then exits, in pressurized form Cp1, through the common connection L0 equipped with valve VL0 (as explained above), before being distributed to each of the nozzles (divergents downstream of the combustion chambers for the deployment of the produced gases) into cooling zones located downstream of the combustion chambers A, A', B, B' to cool them during operation. More specifically, the lines L1-L4 each connect to a cooling line encircling the divergent nozzle downstream of the relevant combustion chamber, for cooling purposes. The heated and pressurized liquid propellant flow Cp1 exiting the cooling zone of each combustion chamber enters the pre-combustion chamber 20 at the inlet via the respective connections 11, 12, 13, 14 already described.The propellant flows supplied to the pre-combustion chamber 20 undergo a combustion reaction with the oxidizer, which is the pressurized propellant E2 from pump 18. The combustion mixture of the two propellants, Cp1 and E2, exits the pre-combustion chamber 20 and reaches the turbine 14, via link L5, to drive the turbine 14 in rotation, as well as pumps 16 and 18. The turbine 14 rotates shaft 14a, which in turn rotates pump 16. Pump 16, in turn, rotates shaft 16a, which in turn rotates pump 18. The rotating pumps 16 and 18 pressurize the incoming propellant flows C1 and C2, respectively. The pressurized mixture E1 not used during combustion in the pre-combustion chamber 20 is distributed to each of the links 14.0, 14.1, and 14.2, as explained above, before being brought to the VE1a and VE1b valves, in each of which the flow is distributed in a controlled manner between the combustion chambers of each pair A, A' and B, B'. In each chamber, combustion occurs between the flows distributed through the different valves, and the resulting thrust is generated. Thus, the flow rate of the propellant flows used to cool the combustion chambers and... The supply to the pre-combustion chamber is controlled by appropriate control of the VLO valve's degree of opening. This also applies to another power source in the power group, such as a gas generator or a turbine from the power group, at the outlet of a heat exchanger.

[0039] It should be noted that the direction of flow of the oxidizing propellant C2 is simpler since it is supplied to the pump 18 to be pressurized before being distributed directly into each of the links 18.0, 18.1 and 18.2, then brought to the valves VE2a and VE2b in each of which the flow is distributed in a controlled manner between the combustion chambers of each pair A, A' and B, B'.

[0040] The circuit in Figure 1 can generate, on command, a propulsion or thrust force.

[0041] In general, the thrust or propulsion force F generated by each chamber is written F = K*Pchamber in which: -K depends on external pressure, -Chamber pressure (chamber pressure) : =K'*(Q ergol E1+Q ergol E2) with K' depending on the properties of the gases and the cross-section of the chamber outlet throat, -Qergol E1 and Qergol E2 are respectively the mass flow rate of propellants E1 and E2, and - F= Q x Ve + (Ps - Pe) x S, where Q is the mass flow rate, Ve is the ejection velocity of the gases at the outlet of the divergent (nozzle), Ps is the static pressure (atmospheric pressure) at the nozzle outlet, Pe is the external pressure and S is the nozzle outlet area.

[0042] As shown in Figure 1, the circuit can generate straight thrust, that is, without any spatial orientation in pitch and yaw: the same thrust FA, FA', FB, FB' is then produced respectively by each of the combustion chambers A, A', B, B'. The valves VE1a and VE2a are adjusted so that the propellant flow rates E1 and E2 of chambers A and A' are such that the propellant flow rate E1 is p 1 1 = 1 2 =Q' in outlet of valve VE1a and the flow rate of propellant E202.1 = 02.2=Q* at the outlet of valve VE2a.

[0043] Thus, in chambers A and A', the thrusts FA and FA' are identical and each includes an identical mass flow rate which is equal to 01.1 + 02.1 = 1.2 + 02.2 = Q'+Q*.

[0044] The same applies to chambers B and B' with valves VE1b and VE2b, which are adjusted so that the propellant flow rates in chambers B and B' are such that 01.1' = 01.2' = Q' at the outlet of valve VE1b and 02.1' = 02.2' = Q* at the outlet of valve VE2b. In chambers B and B', the thrusts FB and FB' are identical and each comprises an identical mass flow rate equal to 01.1' + 02.1 ≈ 01.2' + 02.2' = Q' + Q*, identical to the mass flow rates of thrusts FA and FA'.

[0045] Furthermore, the circuit in Figure 1 can also generate, on command, a spatially oriented thrust (change in direction of thrust): the valves are then set so as to have a different flow rate between the paired chambers A and A', on the one hand, and between the paired chambers B and B', on the other hand.

[0046] In this example, valve VE1a is controlled to distribute a flow rate 01.1 of propellant E1 to chamber A, equal to Q1A + a, and a flow rate 01.2 of propellant E1 to chamber A', equal to Q1A - a, where a is the variation in the E1 propellant flow rate relative to the reference flow rate Q1A (average flow rate). Valve VE2a, on the other hand, is controlled to distribute a flow rate 02.1 of propellant E2 to chamber A, equal to Q2A + P, and a flow rate 02.2 of propellant E2 to chamber A', equal to Q2A - p, where P is the variation in the E2 propellant flow rate relative to the reference flow rate Q2A (average flow rate). This differential flow distribution between chambers A and A' generates a differential thrust: a thrust FA with a mass flow rate equal to (Q1A + a + Q2A + p) is generated in chamber A, while a thrust FA' with a mass flow rate equal to (Q1A - o + Q2A - p) is generated in the associated chamber A' (FA > FA').

[0047] In general, the VE1b and VE2b valves are also controlled to distribute the flow rates of propellants E1 and E2 to be injected into chambers B and B'. Thus, a flow rate of 01.1' = Q1B + y of propellant E1 and a flow rate of 02.1' = Q2B - y of propellant E2 are injected into chamber B, while a flow rate of 1.2' = Q1B + 5 of propellant E1 and a flow rate of 02.2' = Q2B - 5 of propellant E2 are injected into the associated chamber B'. The values ​​y and 5 represent variations from the reference flow rates Q1B and Q2B (average flow rates). In this example, y = 5 = 0, and the thrust FB, FB' generated in each of chambers B and B' is therefore equal.

[0048] The thrust differential (F A > F A' ) associated with the equality of the thrusts FB and FB' thus generates a moment with respect to the axis A1 of figure 2, which allows to obtain a pivoting of the stage of the launcher around this axis in a direction of components [x, y].

[0049] It should be noted that, in a different configuration, the propellant flow rates E1 and E2 in chambers B and B' can be adjusted differently (using valves VE1b and VE2b) so that the thrusts FB and FB' differ from one chamber to the other: the y and 5 values ​​of the flow rate variations Q1B and Q2B are then non-zero. In such a configuration, a moment is generated between chambers B and B' with respect to the median axis between these chambers and passing through the longitudinal axis Z. This moment, coupled with the moment generated between chambers A and A' with respect to the median axis between these chambers and passing through the longitudinal axis Z, thus causes a change in spatial orientation depending on the thrust values.

[0050] For example, the thrust FB can be equal to the thrust FA, while the thrust FB' can be equal to the thrust FA' (FA' < FA). In this case, the launcher stage pivots around the axis A2 in Figure 2.

[0051] Other adjustment configurations and other engine (cycle) arrangements are of course possible.

[0052] Figure 3 schematically illustrates the main components of a launcher control unit configured to command / pilot the various valves VE1a, VE1b', VE2a, VE2b' of distribution and control of the flow rates of ergois E1 and E2 and the VLO control valve of the power group by supplying the pre-chamber 20.

[0053] The launcher's computer (not shown here) generates an overall thrust setpoint Ft (vector) for the launcher stage, defined by a magnitude IFI (absolute value of F) and a direction [x,y] in the plane of Figure 2. The magnitude IFI is the sum of the average thrust IFml and the small thrust variations IFpl around this average. The average thrust IFml, using a control table K (gain), controls the opening of the aforementioned VLO valve, thus regulating the propellant supply to the power unit and modulating the overall engine thrust. The direction of this vector is represented in the plane of Figure 2 by a pair of coordinates (x, y) defining the point of application of the overall thrust. The magnitude of this vector is equal to the sum of the magnitudes of the different thrusts generated by all the combustion chambers: |F| = |FA| + |FA'| + |FB| + |FB'|. These modules are calculated as a function of the direction [x,y],

[0054] This instruction is translated by an AGCV conversion agent which defines, for each pair of matched chambers, A-A' on the one hand, and B-B' on the other, the following instructions for a dedicated MVA or MVB multi-variable controller: - a mixture ratio (MR) instruction to maintain the ratio of the two propellants (mass flow rate of propellant E2 / propellant E1) for each of the chambers, the MR value of which may change during flight, - thrust module instructions |FA|, |FA'| and Ftot [A, A']=[FA|+|FA'| with Ftot [A, A'] constant in accordance with the module |F|, for the dedicated multivariable controller MVA of chambers A and A', - thrust module instructions |FB|, ^FB'^ and Ftot [B, B']=[FB|+|FB'| for the dedicated MVB multivariable controller of chambers B and B'.

[0055] As shown in Figure 3, the multi-variable controller MVA (resp. MVB) also receives, from combustion chambers A (resp. B) and A' (resp. B'), the values ​​PGCA and PGCA' (resp. PGCB and PGCB') which represent the actual pressure values ​​in the chambers and which allow us to determine, for each chamber, the real moduie |FA| or ^FA'^ (^FB^ or |FB'|) of thrust.

[0056] The actual mixing ratio (mass ratio of the two propellants) can be determined by temperature and pressure measurements at the inlet of each chamber (using sensors).

[0057] Based on the thrust module setpoints Ftot, the propellant mixing ratios between the chambers, and the actual pressure values ​​obtained from the chambers, each multi-variable controller (MVA, MVB) calculates the appropriate command to control each valve: VE1a, VE2a for MVA and VE1b, VE2b for MVB. The resulting control law allows for the appropriate control of the valves, thus enabling the controlled distribution of propellant flow rates between the two paired chambers of the same chamber pair (A, A') and (B, B'). Examples of this distribution were given above and explained in relation to Figures 1 and 2.

[0058] Figures 4A and 4B illustrate aspects of the operation of multi-variable controllers for controlling valves VE1a, VE2a, VE1b and VE2b.

[0059] Figure 4A illustrates the operation of the MVA multivariable controller shown in Figure 3 (the controller dedicated to chambers A and A'). This controller receives, as input, a comparison of the setpoints and actual values ​​(sFtot and eRM represent difference signals between the setpoints and actual values ​​for each parameter) for each of the parameters FtotA and RM. As output, it delivers propellant flow rate variation commands for the corresponding valves (VE1a and VE2a), namely ΔVE1A and ΔVE2a. Note that the flow rate variations a and p defined above are such that a is a function of VE1A and β is a function of ΔVE1B. The MVB multivariable controller shown in Figure 3 operates identically.

[0060] Figure 4B, meanwhile, represents the VE2a valve from Figure 1 with, at the outlet, the associated flow rates p2.1 and <t>2.2 intended to supply chambers A and A' respectively and the inlet control flow rate t>2. The other valves VE1a, VE1b and VE2b have the same aspects with respect to the chambers which concern them and with appropriate flow rates.

[0061] It should also be noted that the flow rates 03.1 and 04.1 on Figure 1 are each equal to Q1A and are respectively the flow rate of propellant Cp1 allocated to the inlet and outlet of the cooling line of chamber A. The same is true for all the cooling lines of the other chambers A', B and B' with the flow rate 03.1 and the corresponding flow rate on the respective line I2, I3 and I4.

[0062] It should be noted that everything mentioned above applies to any launcher stage if the piloting dynamics of that launcher allow it (bandwidth) and for an even number of chambers, mainly four chambers, and ranging from two to four pairs of chambers.

[0063] Furthermore, when the launcher has multiple stages, as is the case here, the invention allows, without the need for jacks, the handling of an interstage with limited length, thus avoiding the need for a deployable diverging nose cone. In addition, an upper stage of the launcher (for example, here, the second stage) may include a skirt surrounding the engine. The invention reduces the overall size of the upper stage (by eliminating the use of jacks) and therefore the length of the skirt, consequently reducing the onboard weight.

[0064] Figure 5 illustrates another embodiment of a 10' launcher engine that differs from that of Figure 1 in that only one propellant C1 is used for engine operation. The portion of the diagram in Figure 1 including the pump 18, the fluid lines 18.0, 18.1, 18.2, and the distribution and control valves VE2a, VE2b related to the second propellant C2 (E2) is omitted here.

[0065] The power unit and the fluidic distribution and control circuit of the engine 10' are configured respectively to pressurize a single propellant C1 and to distribute it into each of the 2n combustion chambers, namely here the four chambers A, A', B, B'.

[0066] More specifically, the fluidic distribution and control circuit is configured to distribute the pressurized propellant E1 into the chambers of combustion A, A', B, B' by controlling the distribution of pressurized propellant flow between these fluidly matched combustion chambers so as to inject, for the two pairs of combustion chambers: - a flow rate of propellant 01.1 = Q1A+a in a first combustion chamber A of a first pair and a flow rate of propellant 1.2=01 Aa in a second combustion chamber A' of the first pair, where Q1A is the flow rate of propellant E1 and has a value of variation with respect to this propellant flow rate, - a flow rate of propellant 01.1'=Q1 B+y in a third combustion chamber (B) of a second pair and a flow rate of propellant 01.2'=Q1 By in a fourth combustion chamber B' of the second pair, where Q1 B is the flow rate of propellant E1 and y is a value of variation with respect to this propellant flow rate.

[0067] As shown in Figure 5, the pressurized propellant distribution and control fluid circuit E1 includes fluid lines equipped with distribution and control valves VE1 a, VE1 b which are configured to connect the power unit to the combustion chambers A, A', B, B'.

[0068] More specifically, the fluidic circuit for the distribution and control of the pressurized propellant E1 includes a propellant flow distribution and control valve VE1a, VE1b for each of the fluidly matched combustion chamber pairs A, A', B, B'.

[0069] Everything described above in relation to the two-propellant embodiment of Figures 1 to 4B applies here, except everything concerning the mixing of the two propellants and the distribution of this mixture (RM) which is not relevant in the single-propellant mode.

[0070] It should be noted that regardless of the embodiment (one or two propellants), the cooling of the combustion chambers can be omitted, except in an embodiment with an expansion cycle (heat exchanger).

[0071] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the The general scope of the invention is as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

[0072] It is also evident that all the characteristics described with reference to a process are transposable, alone or in combination, to a device, and conversely, all the characteristics described with reference to a device are transposable, alone or in combination, to a process.< / t>

Claims

Demands [Claim 1] A space launcher engine with at least one stage extending along a longitudinal axis (Z), said engine (10) comprising: -a power unit configured to pressurize at least one propellant to a predetermined pressure, the power unit comprising a pressurization unit (12) for said at least one propellant which is driven by a pre-combustion chamber (20) or by a gas generator or by the outlet of a heat exchanger, -2n combustion chambers (A, A', B, B') of fixed spatial orientation, with n≥2, which are arranged, according to a view taken in a plane perpendicular to the longitudinal axis (Z), two by two diametrically opposite each other on a circumference of a circle centered on the longitudinal axis (Z) and fluidly matched, each of the combustion chambers being able to generate on command a thrust directed along the longitudinal axis (Z), -a fluidic distribution and control circuit for said at least one pressurized propellant which is configured to distribute said at least one pressurized propellant into the 2n combustion chambers (A, A', B, B') by controlling the flow distribution of said at least one pressurized propellant between the fluidly matched combustion chambers so as to inject, for at least one pair of fluidly matched combustion chambers (A, A), a flow Q+α of said at least one propellant (E1) into a first combustion chamber (A) of said at least one pair of fluidly matched combustion chambers (A, A') and a flow Q-α of said at least one propellant (E1) into a second combustion chamber (A') of said at least one pair of fluidly matched combustion chambers (A, A'), where Q is the average flow of said at least one pressurized propellant (E1) from the power group and α a value of variation of this flow. [Claim 2] A space launcher engine according to claim 1, characterized in that the power unit and the fluidic distribution and control circuit are configured respectively to pressurize a single propellant (El) and to distribute it into each of the 2n combustion chambers. [Claim 3] A space launcher engine according to claim 2, characterized in that the fluidic circuit for distributing and controlling the pressurized propellant is configured to distribute the pressurized propellant (E1) into the 2n combustion chambers (A, A', B, B') by controlling the flow distribution of the pressurized propellant between the 2n fluidically paired combustion chambers so as to inject, for at least two pairs of combustion chambers (A, A', B, B'): - a propellant flow rate (E1) Q1A+α in a first combustion chamber (A) of a first pair and a propellant flow rate (E1) Q1A-α in a second combustion chamber (A') of the first pair, where Q1A is the propellant flow rate (E1) and α a value of variation with respect to this propellant flow rate, - a propellant flow rate (E1) Q1B+γ in a third combustion chamber (B) of a second pair and a propellant flow rate (E1) Q1B-γ in a fourth combustion chamber (B') of the second pair, where Q1B is the propellant flow rate (E1) and γ a value of variation with respect to this propellant flow rate. [Claim 4] Space launcher engine according to claim 2 or 3, characterized in that the pressurized propellant distribution and control fluid circuit (E1) comprises fluid lines equipped with distribution and control valves (VE1a, VE1b) which are configured to connect the power unit to the 2n combustion chambers (A, A', B, B'). [Claim 5] A space launcher engine according to claim 1, characterized in that the power unit and the fluidic distribution and control circuit are configured respectively to pressurize two propellants (E1, E2) and to distribute them into each of the 2n combustion chambers. [Claim 6] A space launcher engine according to claim 5, characterized in that the fluidic circuit for the distribution and control of pressurized propellants is configured to distribute the pressurized propellants (E1, E2) in the 2n combustion chambers (A, A', B, B') by controlling the distribution of pressurized propellant flow rates between the fluidly matched combustion chambers so as to inject, for at least two pairs of combustion chambers (A, A', B, B'): - a flow rate of a first propellant (E1) Q1A+α in a first combustion chamber (A) of a first pair and a flow rate of the first propellant Q1A-α in a second combustion chamber (A') of the first pair, -a flow rate of a second propellant (E2) Q2A+β in the first combustion chamber (A) of the first pair and a flow rate of the second propellant (E2) Q2A-β in the second combustion chamber (A') of the first pair, where Q1A and Q2A are respectively the flow rates of the first propellant (E1) and the second propellant (E2) and α, β are the values ​​of variation relative to these flow rates of the first and second propellants, - a flow rate of the first propellant (E1) Q1B+γ in a third combustion chamber (B) of a second pair and a flow rate of the first propellant Q1B-γ in a fourth combustion chamber (B') of the second pair, - a flow rate of the second propellant (E2) Q2B+δ in the third combustion chamber (B) of the second pair and a flow rate of the second propellant (E2) Q2B-δ in the fourth combustion chamber (B') of the second pair, where Q1B and Q2B are respectively the flow rates of the first propellant (E1) and the second propellant (E2) and γ, δ the values ​​of variation with respect to these flow rates of the first and the second propellant. [Claim 7] Space launcher engine according to claim 5 or 6, characterized in that the fluidic circuit for the distribution and control of pressurized propellants comprises fluid lines equipped with distribution and control valves (VE1a, VE2a, VElb, VE2b) which are configured to connect the power unit to the 2n combustion chambers (A, A', B, B'). [Claim 8] Space launcher engine according to any one of claims 4 and 7, characterized in that the distribution and control valves (VEla, VE2a, VElb, VE2b) are electrically or hydraulically controlled. [Claim 9] Space launcher engine according to any one of claims 4, 7 and 8, characterized in that the distribution and control valves (VEla, VE2a, VElb, VE2b) are of the linear or rotary type. [Claim 10] Space launcher engine according to any one of claims 4 and 7 to 9, characterized in that the distribution and control valves (VEla, VE2a, VElb, VE2b) are three-way valves. [Claim 11] A space launcher engine according to any one of claims 4 and 7 to 10, characterized in that it comprises a distribution and control valve for each pressurized propellant intended to supply each pair of fluidly matched combustion chambers. [Claim 12] A space launcher engine according to any one of the preceding claims, characterized in that it comprises at least one control valve (VLO) for controlling a propellant supply flow rate to the power group. [Claim 13] A space launcher with at least one stage extending along a longitudinal axis (Z), the launcher comprising: -a space launcher engine according to one of the preceding claims, -at least one control unit (MVA, MVB) which is configured to control the engine's distribution and control circuit. [Claim 14] Space launcher according to claim 13 and any one of claims 4 and 7 to 11, characterized in that said at least one control unit is configured to control the distribution and control valves (VEla, VE2a, VElb, VE2b) of the engine distribution and control circuit. [Claim 15] Space launcher according to claim 13 or 14 and any one of claims 6 and 7, characterized in that said at least one control unit is configured to control the propellant flow distribution and control valves (VEla, VE2a, VElb, VE2b) according to an overall thrust setpoint and a propellant mixing ratio for the matched combustion chambers. [Claim 16] Method of implementing an engine of a space launcher with at least one stage extending along a longitudinal axis (Z), said engine including: -2n combustion chambers (A, A', B, B') of fixed spatial orientation, with n≥2, which are arranged, according to a view taken in a plane perpendicular to the longitudinal axis (Z), two by two diametrically opposite each other on a circumference of a circle centered on the longitudinal axis (Z) and fluidly matched, each of the combustion chambers being able to generate on command a thrust directed along the longitudinal axis (Z), -a fluidic circuit for the distribution and control of at least one propellant in the 2n combustion chambers, the process comprising: -the pressurization of said at least one propellant (El, E2) to a predetermined pressure from a controlled power unit, -the control of the flow distribution of said at least one pressurized propellant between the fluidly matched combustion chambers (A, A', B, B') so as to inject, for at least one pair of fluidly matched combustion chambers, a flow Q+α of said at least one propellant (E1) into a first combustion chamber (A) of said at least one pair of fluidly matched combustion chambers (A, A') and a flow Q-α of said at least one propellant (E1) into a second combustion chamber (A') of said at least one pair of fluidly matched combustion chambers (A, A'), where Q is the average flow of said at least one pressurized propellant (E1) and α a value of variation of this flow. [Claim 17] Method according to the preceding claim, characterized in that the power group is controlled by controlling a propellant supply rate from a power source.

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