Engine for a space launcher, space launcher and method of implementing such an engine

Abstract

A space launcher engine with at least one stage extending along a longitudinal axis (Z), comprising: - a power unit configured to pressurize at least one propellant to a predetermined pressure, - 2n combustion chambers (A, A', B, B') of fixed spatial orientation, with n ≥ 2, arranged in diametrically opposite pairs and fluidically matched, - a fluidic distribution and control circuit for said at least one propellant configured to distribute said at least one propellant into the chambers (A, A', B, B') by controlling the propellant flow distribution between the matched chambers 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. Figure for the abstract: Fig. 1.

Description

Title of the invention: Engine for a space launcher, space launcher and method for implementing such an engine technical field

[0001] The present exposition relates to a space launcher engine, a space launcher and a method of implementing such a space launcher engine. Previous technique

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

[0003] In a known manner, the orientation of 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] There are known space launcher stage architectures comprising one or more engines equipped with actuators to spatially orient the combustion chambers and thus the overall thrust generated by them. However, actuator activation requires the onboard transport of these heavy components, also necessitates auxiliary components to move them (electronic, hydraulic, pneumatic systems, etc.), and also results in significant size and weight, which can be detrimental to a launcher. 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 comprising: -a power unit configured to pressurize at least one propellant to a predetermined pressure, -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, the combustion chambers being arranged two by two in a diametrically opposite manner and fluidly matched, -a fluidic circuit for the distribution and control of 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 fluidically matched combustion chambers so as to inject, for at least one pair of fluidly matched combustion chambers, a flow rate Q+a of said at least one propellant El in a first combustion chamber of said at least one pair of fluidly matched combustion chambers and a flow rate Qa of said at least one propellant El in a second combustion chamber of said at least one pair of fluidly matched combustion chambers, where Q is the average flow rate of said at least one pressurized propellant El from the power group and a is a value of variation of this flow rate.

[0007] The launcher engine configuration defined above makes it possible to obtain, with combustion chambers of fixed spatial orientation, an overall thrust generated by these chambers whose spatial orientation can vary in the [X, Y] space (perpendicular to the longitudinal axis Z) depending on the distribution of the propellant flow rate supplying the chambers, which are fluidly paired. This configuration does not require actuators to change the orientation of the chambers, which reduces the overall size of the stage, as well as the onboard weight. It should be noted that the average flow rate Q corresponds to the flow rate for a "straight" thrust of the engine, that is to say, a thrust generated solely along the longitudinal axis Z.

[0008] According to other possible characteristics: -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; - The fluidic circuit for the distribution and control of the pressurized propellant is configured to distribute the pressurized propellant El into the 2n combustion chambers by controlling the flow distribution of the pressurized propellant between the 2n fluidically matched combustion chambers so as to inject, for at least two pairs of combustion chambers: - a flow rate QlA+a of propellant El in a first combustion chamber of a first pair and a flow rate QlA-a of propellant El in a second combustion chamber of the first pair, where Q1A is the flow rate of propellant El and has a value of variation with respect to this propellant flow rate, - a Q1B+y flow rate of propellant El in a third combustion chamber of a second pair and a Q1B-y flow rate of propellant El in a fourth combustion chamber of the second pair, where Q1B is the flow rate of propellant El and y is a value of variation with respect to this propellant flow rate; -the fluidic circuit for the distribution and control of the pressurized propellant El 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 the pressurized propellant (El) includes a distribution and control valve (VEla, VElb) for the propellant flow rate for each of the pairs of fluidically matched combustion chambers; -the power unit and the fluidic distribution and control circuit are configured respectively to pressurize two propellants El, 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 QlA+a of a first propellant El in a first combustion chamber A of a first pair and a flow rate of the first propellant QlA-a in a second combustion chamber A' of the first pair, a flow rate Q2A+[3 of a second propellant E2 in the first combustion chamber A of the first pair and a flow rate Q2A-[3 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 El and the second propellant E2 and a, [3 of the values ​​of variation with respect to these flow rates of the first and second propellants, a flow rate Q1B+y of the first propellant El in a third combustion chamber B of a second pair and a flow rate Q1B-y of the first propellant in a fourth combustion chamber B' of the second pair, a flow rate Q1B-y of the second propellant E2 in the third combustion chamber B of the second pair and a flow rate Q2B-Ô of the second propellant E2 in the fourth combustion chamber B' of the second pair, where Q1B and Q2B are respectively the flow rates of the first propellant El and the second propellant E2 and y, ô of 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 the distribution and control of a propellant flow 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 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.

[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 that is configured to control the engine's distribution and control circuit.

[0010] According to 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 for implementing an engine of 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, the combustion chambers being arranged two by two in a diametrically opposite manner and fluidly matched, -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 to a predetermined pressure, -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 El 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 El 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 El and a is a value of variation of this flow.

[0012] The method having the same advantages and characteristics as the launcher engine briefly described above, they will therefore not be repeated here. Brief description of the drawings

[0013] The purpose of this presentation and its advantages will be better understood upon reading the detailed description below of various embodiments given by way of non-limiting examples. This description refers to the attached figure pages, on which:

[0014] [Fig-1] The [Fig. 1] represents a possible embodiment of a space launcher stage engine.

[0015] [Fig.2] Fig.2 represents one possible configuration of the four chambers of combustion of the [Fig.l].

[0016] [Fig.3] The [Fig.3] is a schematic representation of a control unit configured to cause a change in the direction of the thrust of the engine of the [Fig.1].

[0017] [Fig.4A] [Fig.4A] illustrates in more detail an aspect of the operation of the MVA multi-variable controller of [Fig.3].

[0018] [Fig.4B] [Fig.4B] illustrates in detail and enlarged the VE2a valve of [Fig.3],

[0019] [Fig.5] Fig.5 represents another possible embodiment of a space launcher stage engine. Description of the implementation methods

[0020] Figure 1 schematically and partially represents a pump engine 10 of a stage of a one- or multi-stage space launcher. The stage considered here is, for example, the second stage of the launcher.

[0021] The portion of the stage shown in [Fig. 1] comprises a pump engine 10 including a power unit configured to pressurize two propellants, including a fuel Cl (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 thrust operating point of the launcher. The following description relates to a staged combustion engine shown in [Fig. 1] but also applies to other types of engines 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 can be, for example, liquid oxygen (LOX), and the fuel Cl can be either methane (CH4), liquid hydrogen (LH2), or 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 Cl and C2 from tanks (not shown) are identified in the upper part of the diagram. The following discussion will focus on... In general, for pressurized propellants at the output of the power group, El propellant and E2 propellant.

[0022] The engine 10 also includes several combustion chambers, each having a fixed spatial orientation.

[0023] More specifically, the number of chambers is even and equal to 2n, with n > 2. In the present embodiment, the chambers are fluidly paired. In this example of an embodiment illustrated in [Fig. 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.

[0024] Figure 2 shows the geometric arrangement in space [X, Y] (plane perpendicular to the longitudinal axis Z along which the launcher extends) according to 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.

[0025] As shown in [Fig. 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 the circle C: A and A', on the one hand, and B and B', on the other.

[0026] The power group of [Fig.1] includes more particularly a unit 12 for pressurizing the two propellants Cl and C2, which will be described in more detail below.

[0027] In [Fig. 1], the engine 10 also includes a fluidic circuit for the distribution and control of pressurized propellants, denoted E1 and E2 in [Fig. 1], which is configured to distribute the two types of pressurized propellants into 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 and A', on the one hand, and B and B', on the other. This circuit generally comprises: - a plurality of pipes or conduits distributing, on the one hand, the pressurized propellant El 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 VEla, VElb, 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.

[0028] The pressure-generating unit 12 of the power group comprises, for example, a turbine 14, referred to as a 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 Cl 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 which carries a pressurized propellant flow Cpl (at a pressure higher than that of the pressurized propellant flow El), 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 VElb, to the inlets of combustion chambers A, A', B, and B', to which the pressurized propellant E1 is delivered.

[0029] The power unit also includes 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 11, 12, 13, 14 and associated valves (not shown). It should be noted that for a gas generator engine, this would refer to a gas generator and not a pre-combustion chamber. In contrast, for an expansion engine, there is neither a pre-combustion chamber nor a gas generator; the connections 11, 12, 13, 14 supply the turbine 14 directly.

[0030] The prechamber 20 also receives at its inlet the pressurized propellant E2 (oxidizer) flow from pump 18, supplied by valve V5. The prechamber 20 is connected at its outlet to the turbine 14 via a fluid line or connection L5. The function of the prechamber 20 is to generate an enthalpy flow, in this case hot gases from the combustion between the pressurized propellant Cpl from pump 16 and the pressurized propellant E2. The flow thus generated is transmitted to the turbine 14 via connection L5 to drive the turbine 14 in rotation and, consequently, to drive the pumps 16 and 18 in rotation in order to pressurize the propellants Cl and C2. This prechamber 20, together with the pressurization unit 12, together form the engine's power unit.

[0031] The circuit includes, at the outlet of the turbine 14, a common fluidic conduit or link 14.0 distributing the pressurized propellant flow El into two fluidic conduits or links 14.1 and 14.2 which transport fractions of this propellant flow pressurized propellant E2 is distributed to the two valves VE1a and VElb for the distribution and control of the pressurized propellant E2. These valves then distribute the relevant fraction of the pressurized propellant E2 flow to the four combustion chambers A, A', B, and B'. Thus, valve VE1a distributes the propellant E2 in a controlled manner to chambers A and A', and valve VElb distributes the propellant E2 in a controlled manner to chambers B and B'.

[0032] 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 to the two valves VE2a and VE2b for the distribution and control of the pressurized propellant E2. These valves then distribute the relevant fraction of pressurized propellant flow E2 to the four combustion chambers A, A', B, 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'.

[0033] The valves VEla, VElb, VE2a, VE2b are each appropriately electronically controlled by a control unit described later, which distributes the pressurized propellant flows 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.

[0034] In the present embodiment, the valves VEla, VElb, VE2a, VE2b are preferably each of the three-way valves. However, each of these valves can alternatively be replaced by a set of simple two-way valves performing the same function.

[0035] The valves are here electrically controlled but can alternatively be hydraulically controlled. The valves can be of the linear or rotary type.

[0036] It should be noted that a set of simple stop valves VI 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 transients.

[0037] According to the embodiment of [Fig. 1] (staged combustion), the propellant flow Cl is supplied to the pump 16, then exits, in pressurized form Cpl, via the common connection L0 (as explained above), before being distributed to each of the nozzles (divergents downstream of the combustion chambers for the deployment of the produced gases) in cooling zones located downstream of the combustion chambers A, A', B, B' to cool them during their operation. More specifically, the lines L1-L4 each join a cooling line encircling the divergent that is downstream of the combustion chamber concerned, for cooling purposes. The heated and pressurized liquid propellant flow Cpl exiting the cooling zone of each of the combustion chambers joins the The inlet of the combustion chamber 20 is via the respective connections 11, 12, 13, and 14 already described. The flows thus supplied to the combustion chamber 20 undergo a combustion reaction with the oxidant, which is the pressurized propellant E2 from pump 18. The combustion mixture of the two propellants, Cpl and E2, exits the combustion chamber 20 and reaches the turbine 14, via connection 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, which, in turn, rotates shaft 16a, which in turn rotates pump 18. The rotating pumps 16 and 18 pressurize the incoming propellant flows Cl and C2, respectively. The pressurized mixture El not used during combustion in the prechamber 20 is distributed into each of the links 14.0, 14.1 and 14.2, as already explained above, before being brought to the VEla and VElb 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 between the flows / flows distributed through the different valves takes place and the resulting thrust is thus generated.

[0038] It will 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'.

[0039] The circuit of [Fig. 1] can generate, on command, a propulsion or thrust force.

[0040] 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 propellant El + Q propellant E2) with K' depending on the properties of the gases and the cross-section of the chamber outlet throat, -Qergol El and Qergol E2 are respectively the mass flow rate of propellants El and E2, and - F= Q x Ve + (Ps - Pe) x S, where Q is the mass flow rate, Ve the ejection velocity of the gases at the outlet of the divergent (nozzle), Ps the static pressure (atmospheric pressure) at the nozzle outlet, Pe the external pressure and S the nozzle outlet area.

[0041] As shown in [Fig. 1], the circuit can generate straight thrust, i.e. 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 VEla and VE2a are adjusted so that the propellant flow rates El and E2 of chambers A and A' are such that the propellant flow rate El is ¢1 i=Oi 2=Q' at the outlet of the valve VEla and the propellant flow rate E2 ¢2.1= O2.2=Q* at the outlet of the valve VE2a.

[0042] 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 ¢1.1 + ¢2.1 = ¢1.2 + ¢2.2 = Q'+Q::.

[0043] The same applies to chambers B and B' with valves VElb and VE2b, which are adjusted so that the propellant flow rates or fluxes in chambers B and B' are such that ¢L1'=¢L2'=Q' at the outlet of valve VElb and ¢2.1'=¢2.2'=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 ¢1.^+ ¢2.1'= ¢1.2^ ¢2.2^ Q'+Q*, identical to the mass flow rates of thrusts FA and FA'.

[0044] Furthermore, the circuit of [Fig.1] can also generate, on command, a spatially oriented thrust (change of thrust direction): 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.

[0045] In this example, valve VE1 is controlled to distribute a flow rate ¢1.1 of propellant E1 to be injected into chamber A, which is equal to Q1A + a, and a flow rate ¢1.2 of propellant E1 to be injected into chamber A', which is equal to Q1A - a, where a is a value representing the variation in the flow rate of propellant E1 relative to the reference flow rate Q1A (average flow rate). Valve VE2a, on the other hand, is controlled to distribute a flow rate ¢2.1 of propellant E2 to be injected into chamber A, which is equal to Q2A + [3], and a flow rate ¢2.2 of propellant E2 to be injected into chamber A', which is equal to Q2A - [3], where [3] is a value representing the variation in the flow rate of propellant E2 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 + |3) is generated in chamber A, while a thrust FA' with a mass flow rate equal to (Q1A - a + Q2A -13) is generated in the associated chamber A' (FA > FA').

[0046] Generally, the VElb and VE2b valves are also controlled to distribute flow rates of propellants El and E2 to be injected into chambers B and B'. Thus, a flow rate ¢L^=QlB+y of propellant El and a flow rate ¢2.^= =Q2B-y of propellant E2 are injected into chamber B, while a flow rate ¢1.2'=QlB+ô of propellant El and a flow rate ¢2.2'=Q2B-ô of propellant E2 are injected into the associated chamber B'. The values ​​y and ô represent variations from the reference flow rates Q1B and Q2B (average flow rates). In this example, y = ô=0 and the thrust FB, FB' generated in each of chambers B and B' is therefore equal.

[0047] The differential thrust (FA > FA) associated with the equality of the thrusts FB and FB' thus generates a moment with respect to the axis Al of the [Fig.2], which makes it possible to obtain a pivoting of the launcher stage around this axis in a direction of components [x, y].

[0048] It should be noted that, in a different configuration, the propellant flow rates El and E2 in chambers B and B' can be adjusted differently (using valves VElb and VE2b) so that the thrusts FB and FB' are different from one chamber to the other: the y and ô values ​​of the flow rate variations Q1B and Q2B are then non-zero. In such a configuration, a moment is then 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.

[0049] By way of example, the thrust FB may be equal to the thrust FA, while the thrust FB' may be equal to the thrust FA' (FA' <FA). Dans ce cas, l’étage du lanceur pivote autour de l’axe A2 de la [Fig.2].

[0050] Other adjustment configurations and other engine (cycle) arrangements are of course conceivable.

[0051] Fig. 3 schematically illustrates the main components of a launcher control unit configured to command / pilot the various VEla, VElb', VE2a, VE2b' valves for distributing and controlling the flow rates of El and E2 propellants.

[0052] The launcher computer (not shown here) generates an overall thrust setpoint Ft (vector) for the launcher stage, which is defined by a magnitude IFI (absolute value of F) and a direction [x,y] in the plane of [Fig. 2]. The direction of this vector is translated into the plane of [Fig. 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: IFI = IFAI + IFA'I + IFBI + IFB'I. These magnitudes are calculated as a function of the direction [x,y].

[0053] 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 hand, the following instructions for a dedicated multivariable controller MVA or MVB: - a mixture ratio (RM) instruction allowing to maintain the ratio of the two propellants (mass flow rate propellant E2 / propellant El) for each of the chambers, the value of RM being able to evolve during flight, - thrust module instructions IFAI, IFA'I and Ftot [A, A']=IFAI+IFA'I with Ftot [A,A'] constant in accordance with the IFI module, for the dedicated multivariable controller MVA of chambers A and A', - thrust module instructions IFBI, IFB'I and Ftot [B, B']=IFBI+IFB'I for the dedicated multivariable controller MVB of chambers B and B'.

[0054] As shown in [Fig.3], the multi-variable controller MVA (resp. MVB) also receives, from the 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 actual modulus IFAI or IFA'I (IFBI or IFB'I) of thrust.

[0055] 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).

[0056] Based on the thrust module setpoints Ftot and 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 VEla, VE2a for MVA and VElb, VE2b for MVB. The control law thus defined allows the valves to be controlled appropriately, thereby 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 distribution have been given above and explained in relation to Figures 1 and 2.

[0057] Figures 4A and 4B illustrate aspects of the operation of the multi-variable controllers for controlling the VEla, VE2a, VElb and VE2b valves.

[0058] Figure 4A illustrates the operation of the MVA multivariable controller in Figure 3 (the controller dedicated to chambers A and A'), which receives as input a comparison of the setpoints and actual values ​​(eFtot and eRM represent difference signals between the setpoints and actual values ​​for each parameter) for each of the parameters FtotA and RM, and outputs propellant flow rate variation commands for the corresponding valves (VE1a and VE2a), namely AVE1A and AVE2a. It should be noted that the flow rate variations a and 3 defined above are such that a is a function of AVE1A and 3 is a function of AVE1B. The MVB multivariable controller in Figure 3 operates identically.

[0059] Fig. 4B represents the VE2a valve of Fig. 1 with the associated flow rates 2.1 and 2.2 at the outlet, intended to supply chambers A and A' respectively, and the inlet control flow rate 2. The other valves VE1, VElb and VE2b have the same aspects with respect to the chambers they concern and with appropriate flow rates.

[0060] It should also be noted that the flow rates ¢3.1 and ¢4.1 on [Fig.1] are each equal to Q1A and are respectively the flow rate of propellant Cpl 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 ¢3.1 and the corresponding flow rate on the respective line 12, 13 and 14.

[0061] It should be noted that everything mentioned above applies to any launcher stage if the launcher's control dynamics allow it (bandwidth) and for a an even number of rooms, primarily four rooms, and ranging from two to four pairs of rooms.

[0062] Furthermore, when the launcher has several stages, as is the case here, the invention makes it possible, without using jacks, to address 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 makes it possible to reduce the overall size of the upper stage (by eliminating the use of jacks) and therefore the length of the skirt, which consequently reduces the onboard weight.

[0063] Figure 5 illustrates another embodiment of a launcher engine 10' which differs from that of Figure 1 in that only one propellant Cl is used for the operation of the engine. The part 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 in relation to the second propellant C2 (E2) is omitted here.

[0064] The power unit and the fluidic distribution and control circuit of the engine 10' are configured respectively to pressurize a single propellant Cl and to distribute it into each of the 2n combustion chambers, namely here the four chambers A, A', B, B'.

[0065] More specifically, the fluidic distribution and control circuit is configured to distribute the pressurized propellant El into the combustion chambers A, A', B, B' by controlling the flow distribution of the pressurized propellant between these fluidically paired combustion chambers so as to inject, for the two pairs of combustion chambers: - a propellant flow rate ¢1.1= QlA+a in a first combustion chamber A of a first pair and a propellant flow rate O1.2=QlA-a in a second combustion chamber A' of the first pair, where Q1A is the propellant flow rate El and has a value of variation with respect to this propellant flow rate, - a flow rate of propellant Ol.l'=QlB+y in a third combustion chamber (B) of a second pair and a flow rate of propellant O1.2'=QlB-y in a fourth combustion chamber B' of the second pair, where Q1B is the flow rate of propellant El and y is a value of variation with respect to this propellant flow rate.

[0066] As shown in [Fig.5], the pressurized propellant distribution and control fluid circuit El includes fluid lines equipped with distribution and control valves VE la, VElb which are configured to connect the power unit to the combustion chambers A, A', B, B'.

[0067] More specifically, the fluidic circuit for distributing and controlling the pressurized propellant El includes a distribution and control valve VEla, VElb with a flow rate propellant for each of the pairs of fluidly matched combustion chambers A, A', B, B'.

[0068] 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 embodiment.

[0069] It should be noted that regardless of the embodiment (one or two propellants), the cooling of the combustion chambers can be omitted.

[0070] 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 general scope of the invention 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 a restrictive sense.

[0071] 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.

Claims

Demands

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, - 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), the combustion chambers (A, A', B, B') being arranged in diametrically opposite pairs and fluidically matched, - a fluidic circuit for the distribution and control of 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 rate Q+a of said at least one propellant (El) in a first combustion chamber (A) of said at least one pair of fluidly matched combustion chambers (A, A') and a flow rate Qa of said at least one propellant (El) in a second combustion chamber (A') of said at least one pair of fluidly matched combustion chambers (A, A'), where Q is the average flow rate of said at least one pressurized propellant (El) from the power group and a is a value of the variation of this flow rate.

2. A space launcher engine according to claim 1, characterized in that the power group 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.

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 (El) 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 (El) QlA+a in a first combustion chamber (A) of a first pair and a propellant flow rate (El) QlA-a in a second combustion chamber (A') of the first pair, where Q1A is the propellant flow rate (El) and a value of variation with respect to this propellant flow rate, - a propellant flow rate (El) QlB+y in a third combustion chamber (B) of a second pair and a propellant flow rate (El) QlB-y in a fourth combustion chamber (B') of the second pair, where Q1B is the propellant flow rate (El) and y a value of variation with respect to this propellant flow rate.

4. Space launcher engine according to claim 2 or 3, characterized in that the pressurized propellant distribution and control fluid circuit (El) comprises fluid lines equipped with distribution and control valves (VEla, VElb) which are configured to connect the power group to the 2n combustion chambers (A, A', B, B').

5. A space launcher engine according to claim 1, characterized in that the power group and the fluidic distribution and control circuit are configured respectively to pressurize two propellants (El, E2) and to distribute them into each of the 2n combustion chambers.

6. A space launcher engine according to claim 5, characterized in that the fluidic circuit for distributing and controlling the pressurized propellants is configured to distribute the pressurized propellants (E1, E2) into the 2n combustion chambers (A, A', B, B') by controlling the distribution of the pressurized propellant flow rates between the fluidically paired 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+a into a first combustion chamber (A) of a first pair and a flow rate of the first propellant Q1A-a into a second combustion chamber (A') of the first pair, - a flow rate of a second propellant (E2) Q2A+[3] into the first combustion chamber (A) of the first pair and a flow rate of the second propellant (E2) Q2A-[3] into the second combustion chamber (A') of the first pair, where Q1A and Q2A are respectively the flow rates of the first propellant (El) and the second propellant (E2) and a, [3 of the values ​​of variation with respect to these flow rates of the first and second propellant, - a flow rate of the first propellant (El) Q1B+y in a third combustion chamber (B) of a second pair and a flow rate of the first propellant Q1B-y in a fourth combustion chamber (B') of the second pair, - a flow rate of the second propellant (E2) Q2B+Ô in the third combustion chamber (B) of the second pair and a flow rate of the second propellant (E2) Q2B-Ô in the fourth combustion chamber (B') of the second pair, where Q1B and Q2B are respectively the flow rates of the first propellant (El) and the second propellant (E2) and y, ô of the values ​​of variation with respect to these flow rates of the first and second propellant.

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').

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.

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.

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.

11. A space launcher engine according to any one of the preceding claims, characterized in that the power group comprises 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.

12. 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.

13. Space launcher according to claim 12 and any one of claims 4 and 7 to 10, 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.

14. Space launcher according to claim 13 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.

15. A method for implementing an engine for a space launcher with at least one stage extending along a longitudinal axis (Z), said engine comprising: - 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), the combustion chambers (A, A', B, B') being arranged in diametrically opposite pairs and fluidically matched, - a fluidic circuit for distributing and controlling at least one propellant in the 2n combustion chambers, the method comprising: - pressurizing said at least one propellant (E1, E2) to a predetermined pressure, - controlling the flow distribution of said at least one pressurized propellant between the fluidically matched combustion chambers (A, A', B, B') so as to inject, for at least one pair of fluidly matched combustion chambers,a flow rate Q+a of said at least one propellant (El) in a first combustion chamber (A) of said at least one pair of fluidly matched combustion chambers (A, A') and a flow rate Qa of said at least one propellant (El) in a second combustion chamber (A') of said at least one pair of matched combustion chambers, fluidly (A, A'), where Q is the average flow rate of said at least one pressurized propellant (El) and has a value of variation of this flow rate.

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