Rocket engine, propulsion system and method for vehicles, in particular for space vehicles

Abstract

A rocket engine has a convergent-divergent nozzle (12) having a minimum passage section (Smin) configured to guide a propellant flow comprising fuel and oxidant and combustion product in a given direction (D1) and extending along a given axis (A); a reaction chamber (5) arranged upstream of the convergent- divergent nozzle (12) and in which the fuel and oxidant are subject to a chemical reaction that generates energy in the form of heat, the reaction chamber (5) comprising in series along the axis an injection and atomization zone (Z1), a quick combustion zone (Z2), and a streamtube combustion zone (Z3); an ionization device (6), which is configured to ionize the propellant and comprises at least a first ionization station (15) arranged within the reaction chamber (5) for plasma assisted combustion in the quick combustion zone (Z2) and ionizing the fuel and the oxidant and the combustion products; and a plasma propulsion engine accelerator (7) downstream of the convergent-divergent nozzle (12).

Description

[0001] "ROCKET ENGINE, PROPULSION SYSTEM AND METHOD FOR VEHICLES, IN PARTICULAR FOR SPACE VEHICLES"

[0002] Cross-Reference To Related Applications

[0003] This patent application claims priority from European patent application no. 24425004.9 filed on February 23, 2024, the entire disclosure of which is incorporated herein by reference.

[0004] Technical Field

[0005] The present invention concerns a rocket engine, a propulsion system and method for vehicles, in particular for space vehicles .

[0006] Rocket engines exploit the principle of action and reaction to produce a thrust transforming the kinetic energy transferred to the propellant. In practice, the propellant is accelerated and subsequently ejected in an opposite direction to the desired thrust direction.

[0007] The thrust F generated by a rocket engine is determined by the product of the ejection velocity of the propellant from the engine veand the flow of propellant ejected in which F is the propulsion force measured in Newton, is the mass ejected from the reactor expressed in kg / s, and is the velocity expressed in m / s .

[0008] The efficiency of a rocket engine is commonly evaluated by means of the specific impulse Isp, which represents the inverse of the specific consumption of propellant cs. This parameter is directly proportional to the propellant ejection velocity. Referring to the equation in the preceding paragraph, the specific impulse provides an indication of the efficiency with which a rocket engine exploits its propellant flow to obtain a thrust .

[0009] The specific impulse Ispis measured in seconds [s] ; where go is the gravitational constant at sea level, equivalent to 9.81 metres per square second [m / s2] , F is the propulsive thrust, measured in Newton [N] , cs is the specific consumption of propellant in terms of thrust, expressed in kilograms per Newton-second [kg / Ns] .

[0010] State of the Art

[0011] As described in the document US 2009 / 229,240 Al, rocket engines are grouped into different categories according to propulsion type. The two categories of rocket engines most widely used are chemical engines and electric engines. The so-called chemical rocket engines are based on a chemical reaction of the propellant and are characterized by the high thrusts generated due to the capacity to eject large quantities of mass generated by exothermic processes. However, the efficiency of this type of rocket engine is limited by the thermodynamic processes involved .

[0012] Electric engines are based on the use of electromagnetic fields for acceleration of the propellant; hence they have low consumption, but they are also characterized by relatively low thrusts, limited by the high energy demand per unit of mass processed .

[0013] The document US 2009 / 229,240 Al describes the production of a hybrid chemical-electric rocket engine aimed at obtaining a flexible technology .

[0014] Subject of the Present Invention

[0015] The obj ect of the present invention is to produce a rocket engine able to further improve the performances of rocket engines in the known art .

[0016] In accordance with the present invention, a rocket engine for vehicles , in particular for space vehicles , the rocket engine comprising :

[0017] - a convergent-divergent noz zle having a minimum passage section configured to guide a propellant flow comprising fuel and oxidant and combustion product in a given direction and extending along a given axis ;

[0018] - a reaction chamber arranged upstream of the convergent- divergent noz zle and in which the fuel and oxidant are subj ect to a chemical reaction that generates energy in the form of heat , the reaction chamber comprising in series along the axis an inj ection and atomi zation zone , a quick combustion zone , and a streamtube combustion zone ;

[0019] - an ioni zation device , which is configured to ioni ze the propellant and comprises at least a first ioni zation station arranged within the reaction chamber for plasma assisted combustion in the quick combustion zone and ionizing the fuel and the oxidant and the combustion products ; and

[0020] - a plasma propulsion engine accelerator downstream of the convergent-divergent noz zle .

[0021] The present invention introduces a first ioni zation station within the reaction chamber in the quick combustion zone . This configuration allows the properties of the chemical reaction process to be locally improved, at least partially beginning to generate the plasma which is subsequently accelerated via the plasma propulsion engine accelerator at the outlet of the convergent-divergent noz zle . The interaction between the plasma and the chemical reaction zone , in particular the quick combustion zone , is complex and di f ficult to predict . The phenomenon of 'plasma assisted combustion' improves the stability of the flame , catalyses the combustion processes , allows more stable and rapid combustion reactions to be obtained, also in adverse conditions , allows the use of leaner mixtures , and reduces the polluting products of unburnt material . Plasma assi sted combustion takes place when a device generates an electric field near a spatial region in which a combustion reaction occurs . In the case of suf ficiently intense electromagnetic interactions , the chemical species present in said region absorb energy, first kinetically, then electronically, reaching first a level such as to allow the dissociation thereof and subsequently the ej ection of electrons (namely, achieving ioni zation) . Said chemical species therefore constitute a plasma, within which constant interactions occur between free electrons and ions ( sometimes free radicals ) . These species are characteri zed by a strong chemical reactivity, which therefore allows the combustion process to be started and sustained, also in conditions in which it would not be able to sustain itsel f in a stable manner . The ioni zed propellant in the plasma region mentioned continues its motion along the engine axis , recombining and partially losing its plasma state , but a fraction of it maintains suf ficient thermal energy to remain ioni zed . Said event indicates that the propellant emitted from the engine is partially composed of ioni zed particles , and can therefore be influenced by an external electromagnetic field .

[0022] For a plasma to be sustained, the electromagnetic interactions between its particles must prevail over those due to the kinetic collisions , a requirement that can be di f ficult to meet in high pressure environments . This is due to the fact that , at high pressures , the frequency of the collisions between particles increases , thus favouring cooling and recombination of the ions and the electrons , leading to decay of the plasma . Therefore , maintaining a stable plasma requires careful balancing of the operating conditions to prevent premature decay thereof , especially after leaving the primary ioni zation zone .

[0023] In particular, the ioni zation device comprises a second ioni zation station within the reaction chamber in the streamtube combustion zone in close proximity of the minimum passage section and arranged between the first ioni zation station and said minimum passage section .

[0024] This further ioni zation contributes to generating further plasma .

[0025] In the described configuration the second ioni zation station can be of the same nature as the first or di f ferent . Its function is mainly to support the first station, relatively influencing the performance of the propellant as described in the present invention . In di f ferent application variants , the second station can be active , contributing to the improvement of the plasma generation region, or inactive or eventually omitted

[0026] In particular, the plasma propulsion engine accelerator comprising electric field generator, which is arranged at the outlet of the convergent-divergent noz zle and comprises acceleration electrodes .

[0027] The plasma propulsion engine accelerator may comprise in addition to the magnetic field generator, which is arranged around the electric field generator . In practice , the electric field generator is configured to generate an electric field parallel to the advancement direction of the propellant while the magnetic field generator is configured to generate a radial magnetic field perpendicular to the advancement direction of the propellant . Both the magnetic field and the electric field cause an acceleration of the propellant as it is ej ected from the rocket engine .

[0028] The electric field generator must guarantee a linear electric field, aligned with the engine axis , while the magnetic field generator must guarantee a magnetic field that acts as a selector, determining di f ferent traj ectories for ioni zed particles having opposite charge . The electric field imposes an acceleration on the particles of ioni zed propellant , allowing direct control over the velocity of the outgoing propellant . Since the speci fic impulse is defined by this relation, controlling the velocity of the particles means directly influencing the engine performance . In this way the propellant can be accelerated in variable modulus (by introducing energy) or also decelerated (by absorbing energy from the device ) . The magnetic field, on the other hand, regulates the action of the electric field : the plasma coming out of the combustion chamber is almost neutral , composed of positively and negatively charged particles . Simply accelerating the flow would produce a null ef fect , since the contribution of quantity of motion of a group of particles would be annulled by the opposite one of the opposite charge group . Magnetic fields designed to rapidly divert the electrons , without signi ficantly influencing the traj ectories of the ions , remove the negative contribution of the electrons , al lowing only the positive ioni zed particles to determine the flow properties .

[0029] In the case in point , the electric field generator comprises acceleration electrodes .

[0030] The electric field generator comprises at least two acceleration electrodes , designed to be able to interact directly with the environment downstream of the engine outlet section . By means of an electrical energy distribution system, they are supplied with direct current to establish an electric potential gradient in the spatial region between them .

[0031] In accordance with an embodiment of the present invention, the ioni zation device comprises a third ioni zation station at an outlet of the convergent-divergent noz zle . In fact, the thermodynamic expansion process characteristic of a convergent- divergent noz zle involves another region which is highly critical for survival of the plasma . The rapid conversion of energy from thermal to kinetic undergone by the propellant during expansion, in addition to its consequent density reduction, configure a condition in which the remaining plasma is partly deprived of the energy necessary to maintain its state . To actuate the electronic acceleration process for plasma propulsion engines previously described, the addition of a further ioni zation station is expedient , in order to exploit and ampli fy the plasma fraction that has survived the expansion and make the propellant even more subj ect to influence by external actuations .

[0032] According to an alternative embodiment , the function of the third ioni zation station can be carried out by the electric field generator itsel f . The latter, for certain operating parameters , can be able to favour the formation of additional plasma, for example by firing a non-equilibrium plasma discharge . The three ioni zation stations can be selectively activated in parallel in an alternate manner and other combinations in a wide range of modulations excite an intense electric field that strikes a plasma arc in the spatial region between said ioni zation stations .

[0033] According to further preferred embodiment , ioni zation stations can generate plasma independently and through mechanisms other than the electric arc . By way of example and not limited to , this can occur through non-equilibrium plasma discharges .

[0034] A further obj ect of the present invention is to provide a propulsion system for vehicles , in particular for space vehicles that is able to improve the performance of the known systems .

[0035] In accordance with the present invention i s reali zed a propulsion system is provided for vehicles , in particular space vehicles , the system comprising a rocket engine for vehicles as disclosed above and at least two tanks configured to contain respectively the fuel and oxidant and selectively supply the reaction chamber .

[0036] In this way, the system is able to improve the performance of the speci fic impulse and reduce fuel consumption .

[0037] In particular, the system comprises an electrical energy source to power the ioni zation device and the accelerator for plasma propulsion engines .

[0038] The energy source can be of renewable type such as , for example , solar energy and storage systems . The type of electrical energy source depends on the mission profile . For launch missions , where high power is required, the use of a high ef ficiency turbogenerator is preferable . For orbital manoeuvres , on the other hand, the use of solar panels or radioisotope generators is more appropriate .

[0039] In further detail , the system provides an electrical energy distribution bus to connect the electrical energy source to the ioni zation device and to the plasma propulsion engine accelerator . In practice , the bus is connected to each ioni zation station .

[0040] In particular, the system provides a power control unit configured to modulate the electrical energy supplied to the ioni zation device and to the plasma propulsion engine accelerator and modulate each ioni zation station independently of the other ioni zation stations .

[0041] The power control unit is also configured to trans fer energy to any storage units when the engine is in deceleration mode .

[0042] A further obj ect of the present invention is to provide a propulsion method for vehicles , in particular for space vehicles , which is free from the drawbacks of the known art .

[0043] In accordance with the present invention, a propulsion method in particular for space vehicles , the method comprising :

[0044] - supplying a propellant comprising fuel and oxidant to a reaction chamber of a convergent-divergent noz z le having a minimum passage section ( Smin) and comprising in series along the axis upstream of the minimum passage section an inj ection and atomi zation zone , a quick combustion zone , and a streamtube combustion zone ;

[0045] - ioni zing at least partly the fuel and the oxidant within the reaction chamber in the quick combustion zone to generate plasma by means of a first ioni zation station so as to generate a plasma assisted combustion of the fuel and the oxidant ;

[0046] - guiding a flow of at least partly ioni zed propellant in a given direction through a convergent-divergent noz zle having a minimum passage section ; and accelerating the at least partly ioni zed flow of propellant by means of a plasma propulsion engine accelerator downstream of the convergent-divergent noz zle .

[0047] In this way it is possible to improve the speci f ic impulse and reduce the consumption of propellant while obtaining the same performance .

[0048] The first ioni zation station generates non-equilibrium plasma . In this configuration, the system is able to start the chemical reaction and sustain plasma generation with j ust one ioni zation station .

[0049] What has been introduced by this new invention is a propulsion method capable of adapting the concept of multimodal electrochemical propulsion to an engine based on the phenomenon of plasma-assisted combustion . By managing the flow of plasma created in the combustion chamber, after it has participated in the chemical combustion reactions , it is possible to exploit its electromagnetic properties to implement a second acceleration . This will add a component to the thrust generated by the engine ; where a first gas dynamic contribution typically reaches outflow speeds between 1900 and 4900 m / s , the contribution of the plasma phase is in the order of magnitude of 10000 m / s and can reach 100000 m / s . The thrust of the propulsion system will therefore be given by the linear combination of the two contributions based on the fraction of ioni zed mass f according to the equation where the parameters F and rh are the same as those described in the introduction, f is a dimensionless parameter indicating the ratio between the fraction of ioni zed and non-ioni zed mass in the ef flux, the velocities v_e , gas and v_e , plasma respectively the ef flux velocity of the gas phase and ioni zed phase in m / s . The fact that within the same engine the plasma is used both as a catalyst for the chemical reaction (plasma assisted combustion) and as a second propulsive method (plasma acceleration) means that there is a greater energy advantage than simply overlapping the two single ef fects .

[0050] Brief Description of the Figures

[0051] Further characteristics and advantages of the present invention will appear clear from the following description of non-limiting embodiment examples thereof , with reference to the Figures of the attached drawings , in which :

[0052] - Figure 1 is a schematic view, with parts removed for clarity, of a propulsion system produced in accordance with the present invention of a rocket engine for vehicles produced in accordance with the present invention;

[0053] - Figure 2 is a schematic view, with parts removed for clarity, of a rocket engine for vehicles produced in accordance with the present invention; and

[0054] - Figure 3 is a schematic view, with parts removed for clarity and on an enlarged scale , of a detail of the rocket engine of Figure 2 .

[0055] Preferred Embodiment of the Invention

[0056] With reference to Figure 1 , the number 1 indicates overall a propulsion system for vehicles , in particular for space vehicles , not shown in the attached Figures .

[0057] The system 1 comprises a rocket engine 2 , at two tanks 3A and 3B, which are configured to contain propellants and feed the propellants to the rocket engine 2 where the propellants react and develops energy in the form of heat ; and a source of electrical energy 4 to control the rocket engine 2 .

[0058] The propellants comprise a fuel , for example ethanol , hydrogen, rocket grade kerosene RP- 1 and others , contained in tank 3A and an oxidant , for example oxygen, nitrous oxide , nitrogen tetroxide and others contained in tank 3B .

[0059] The rocket engine 2 comprises a reaction chamber 5 in which the propellant , namely the fuel and the oxidant react and develop energy in the form of heat , an ioni zation device 6 to ioni ze the propellant and form plasma, and a plasma propulsion engine accelerator 7 , which comprises an electric field generator 8 and a magnetic field generator 9 .

[0060] The system 1 comprises an electrical energy distribution bus 10 to connect the electrical energy source 4 to the ioni zation device 6 and to the plasma propulsion engine accelerator 7 , and a power control unit 11 to modulate the electrical energy supplied to the ioni zation device 6 and to the plasma propulsion engine accelerator 7 .

[0061] With reference to Figure 2 , the rocket engine 2 extends along an axis A and comprises , downstream of the reaction chamber 5 , a convergent-divergent noz zle 12 configured to guide a propellant flow in a direction DI parallel to the axis A and having a minimum passage section Smin that ideally divides the noz zle 12 into a portion upstream 13 and a portion downstream 14 of the minimum passage section Smin, commonly called throat in the speci fic reference sector . The combustion chamber 5 is ideally divided into four zones Zl, Z2, Z3, Z4 distributed in succession along the axis in accordance with the prevalent physical events occurring in these zones Zl, Z2, Z3, Z4.

[0062] Namely, zone Zl identifies an atomization zone where the fuel and oxidant are atomized, zone Z2 directly downstream of the zone Zl identifies a quick combustion zone characterized by a turbulent motion of the fuel, the oxidant and the combustion products, and a streamtube combustion zone Z3 downstream of zone Z2 and upstream of the minimum passage section Smin is characterized by a residual reaction and a less turbulent motion of fuel and oxidant and combustion products.

[0063] Zone Z4 arranged downstream of the minimum passage section Smin is characterized by a supersonic flow.

[0064] The ionization device 6 comprises three ionization stations 15, 16 and 17 distributed along the combustion chamber 5. The station 15 is arranged in zone Z2 characterized by a quick reaction of fuel and oxidant, the station 16 is arranged in zone Z3 directly upstream of the minimum passage section Smin where the reaction is over, and ionization station 17 is arranged at the outlet of the nozzle 12.

[0065] The plasma propulsion engine accelerator 7 is arranged at the outlet of the nozzle 12 and, in particular, the electric field generator 8 is arranged downstream of the ionization station 17, and the magnetic field generator 9 is arranged around the electric field generator 8 at the outlet of the nozzle 12.

[0066] The electric field generator 8 comprises acceleration electrodes while the magnetic field generator 9 comprises opposite solenoids at the ends of a diameter of the outlet section. With reference to Figure 3 , the electric field generator 8 generates an electric field E in the direction DI while the magnetic field generator 9 generates a magnetic field B in a radial direction transversely to the direction DI to accelerate the plasma and impart a velocity V to the plasma .

[0067] With reference to Fig . 3 , rocket engine 2 comprises upstream of zone 1 a mani fold 18 for the oxidant and a mani fold 19 for the fuel a plurality of inj ection noz zles 20 for the fuel distributed around axis A along annular path and a plurality of inj ection noz zles 21 for the fuel distributed around axis A along annular path internal to the distribution annular path of the inj ection noz zles 20 .

[0068] More in detail , ioni zation station 6 comprises a plurality of electrodes 22 arranged in the zone Z2 distributed along an annular path around axis A. Each electrode 22 extends radially through the wall of the combustion chamber 5 and is aligned to an opposite electrode 22 . Electrodes 22 of each couple of opposite electrodes 22 are configured to generate an electrical discharge and are controlled either in a fixed polarity or in an alternating manner . The alternating manner can span from the emission of microwaves , to radiofrequency waves or to nanosecond pulsed discharges according to the modulation of the power controller 11 .

[0069] The configuration of the ioni zation station 16 and 17 could be the same as that of ioni zation station 15 .

[0070] In use , the fuel and the oxidant are inj ected onto the noz zle 12 and atomi zed in zone Z l , then the reaction occurs in Z2 , where the ioni zation of the fuel and the oxidant promotes the chemical reaction by lowering the activation energy thereof , in accordance with the so-called plasma assisted combustion . Plasma assisted combustion improves the ef ficiency of combustion and reduces the energy gap to start the combustion .

[0071] The combustion products are still ioni zed, however, depending on di f ferent circumstances like the si ze of the combustion chamber 5 , density of the fluid and flame temperature , the ioni zation of the chemical species may decay considerably along axis A in direction DI prior to reaching a plasma motor accelerator 7 .

[0072] In this case , it would be necessary to ioni ze again the combustion products using the second ioni zation station 16 or even the third ioni zation station 17 in order to optimi ze the ef ficiency of rocket engine 2 .

[0073] In such a case , the energy required for the ioni zation of the chemical species deriving from the plasma assisted combustion is lower than in Z2 because the fluid flow still carries a portion of excited and ioni zed species , these increasing overall the energy level and lowering the gap to ioni zation .

[0074] The actuation of the ioni zation station 15 , 16 and 17 and the accelerator 7 is performed through the power control unit 11 ( Figure 1 ) .

[0075] Propulsion system 1 can operate in di f ferent manner :

[0076] - combined mode comprising reaction of fuel and generation and acceleration of plasma as described above ;

[0077] - thermal mode comprising reaction of fuel and oxidant without with plasma generation and subsequent acceleration of the same ; and

[0078] - plasma mode comprising inj ection of the fuel or the oxidant and ioni zation and acceleration of engine .

[0079] Lastly, it is evident that the present invention comprises further variations of the embodiments described included in the protective scope of the attached claims.

[0080] 5

Claims

CLAIMS1. A rocket engine for vehicles, in particular for space vehicles, the rocket engine (2) comprising: a convergent-divergent nozzle (12) having a minimum passage section (Smin) configured to guide a propellant flow comprising fuel and oxidant and combustion product in a given direction (DI) and extending along a given axis (A) ; a reaction chamber (5) arranged upstream of the convergent-divergent nozzle (12) and in which the fuel and oxidant are subject to a chemical reaction that generates energy in the form of heat, the reaction chamber (5) comprising in series along the axis an injection and atomization zone (Zl) , a quick combustion zone (Z2) , and a streamtube combustion zone (Z3) ;- an ionization device (6) , which is configured to ionize the propellant and comprises at least a first ionization station (15) arranged within the reaction chamber (5) for plasma assisted combustion in the quick combustion zone (Z2) and ionizing the fuel and the oxidant and the combustion products; and- a plasma propulsion engine accelerator (7) downstream of the convergent-divergent nozzle (12) .

2. The rocket engine as claimed in claim 1, wherein the ionization device (6) comprises a second ionization station (16) within the reaction chamber (5) in the streamtube combustion zone (Z3) in close proximity of the minimum passage section ( Smin) and arranged between the first ionization station (15) and said minimum passage section ( Smin) .

3. The rocket engine as claimed in claim 1 or 2, wherein the plasma propulsion engine accelerator (7) comprising electricfield generator (8) , which is arranged at the outlet of the convergent-divergent nozzle (12) and comprises acceleration electrodes .

4. The rocket engine as claimed in claim 3, wherein the plasma propulsion engine accelerator (7) comprising a magnetic field generator (9) , which is arranged around the electric field generator ( 8 ) .

5. The rocket engine as claimed in any one of the foregoing claims, wherein the ionization device (6) comprises a third ionization station (17) at the outlet of the convergent- divergent nozzle (12) .

6. A propulsion system for vehicles, in particular space vehicles, the system (1) comprising:- a rocket engine (2) for vehicles as claimed in any of the preceding claims; and- at least two tanks (3A, 3B) configured to contain respectively the fuel and oxidant and selectively supply the reaction chamber (5) .

7. The system as claimed in claim 6, comprising an electrical energy source (4) to power the ionization device (6) and the plasma propulsion engine accelerator (7) and an electrical energy distribution bus (10) to connect the electrical energy source (4) to the ionization device (6) and to the plasma propulsion engine accelerator (7) .

8. The system as claimed in claim 7, comprising a power control unit (11) configured to modulate the electrical energy supplied to the ionization device (6) and to the plasma propulsion engine accelerator (7) and controlling the injection of the fuel andthe oxidant.

9. A propulsion method for vehicles, in particular for space vehicles, the method comprising:- supplying a propellant comprising fuel and oxidant to a reaction chamber of a convergent-divergent nozzle (12) having a minimum passage section (Smin) and comprising in series along the axis upstream of the minimum passage section (Smin) an injection and atomization zone (Zl) , a quick combustion zone (Z2) , and a streamtube combustion zone (Z3) ;- ionizing at least partly the fuel and the oxidant within the reaction chamber (5) in the quick combustion zone (Z2) to generate plasma by means of a first ionization station (15) so as to generate a plasma assisted combustion of the fuel and the oxidant ;- guiding a flow of at least partly ionized propellant in a given direction through a convergent-divergent nozzle (12) having a minimum passage section (Smin) ; and accelerating the at least partly ionized flow of propellant by means of a plasma propulsion engine accelerator (7) downstream of the convergent-divergent nozzle (12) .

10. The method as claimed in claim 9 comprising ionizing the propellant by means of a second ionization station (16) arranged within the reaction chamber (5) streamtube combustion zone (Z3) between the first ionization station (15) and said minimum passage section (Smin) at said minimum passage section (Smin) .

11. The method as claimed in claim 9 or 10 comprising ionizing the propellant by means of a third ionization station (17) arranged at the outlet of the convergent-divergent nozzle (12) .

12. The method as claimed in any one of the claims from 9 to 11comprising supplying the ionization device (6) and the plasma propulsion engine accelerator (7) with electrical energy.

13. The method as claimed in claim 12 comprising modulating the electrical energy supplied to the ionization device (6) and to the plasma propulsion engine accelerator (7) .

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