Method for operating a rocket propulsion system and rocket propulsion system

Abstract

A method for operating a rocket propulsion system comprises the steps of supplying oxygen to a combustion chamber, supplying hydrogen to the combustion chamber and combusting the oxygen-hydrogen mixture in the combustion chamber. The rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.

Claims

1-15. (canceled)

16. A method for operating a rocket propulsion system, which comprises: supplying oxygen to a combustion chamber, supplying hydrogen to the combustion chamber, and combusting the oxygen-hydrogen mixture in the combustion chamber, wherein the rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.

17. The method according to claim 16, wherein the first mass mixing ratio is a sub-stoichiometric mass mixing ratio of oxygen to hydrogen, and wherein the second mass mixing ratio is a super-stoichiometric mass mixing ratio of oxygen to hydrogen.

18. The method according to claim 17, wherein the first mass mixing ratio is less than 2.

19. The method according to claim 17, wherein the second mass mixing ration is greater than or equal to 50.

20. The method according to claim 17, wherein in the first operating mode of the rocket propulsion system, a first hydrogen mass flow is supplied to the combustion chamber that is greater than a second hydrogen mass flow, which is supplied to the combustion chamber in the second operating mode of the rocket propulsion system.

21. The method according to claim 16, wherein the supply of hydrogen to the combustion chamber is interrupted in the second operating mode of the rocket propulsion system.

22. The method according to claim 16, wherein in the first operating mode of the rocket propulsion system, a first oxygen mass flow is supplied to the combustion chamber that is smaller than a second oxygen mass flow which is supplied to the combustion chamber in the second operating mode of the rocket propulsion system.

23. The method according to claim 16, wherein at least a portion of the oxygen supplied to the combustion chamber and at least a portion of the hydrogen supplied to the combustion chamber are conducted into a catalyst chamber and the combustion of the oxygen-hydrogen mixture is initiated in the catalyst chamber, and wherein at least one of: a flashback arrestor is arranged in the region of an entrance area to the catalyst chamber; the oxygen conducted into the catalyst chamber and the hydrogen conducted into the catalyst chamber are premixed in a premixing chamber of the catalyst chamber prior to initiating the combustion of the oxygen-hydrogen mixture; oxygen is supplied to the catalyst chamber via an oxygen supply opening, which is formed in a wall of the catalyst chamber facing an oxygen supply duct; or hydrogen is supplied to the catalyst chamber via a hydrogen supply duct opening into the catalyst chamber or via a hydrogen supply opening which is formed in a wall of the catalyst chamber facing a hydrogen supply duct.

24. The method according to claim 23, wherein the catalyst chamber is enclosed, at least in sections, by a cooling duct arranged between an outer surface of the catalyst chamber and an inner surface of the combustion chamber, which duct opens into a combustion section of the combustion chamber arranged downstream of an exit area of the catalyst chamber and through which duct oxygen supplied to the combustion chamber or hydrogen supplied to the combustion chamber flows, wherein at least one of: oxygen flows through the cooling duct only in the second operating mode of the rocket propulsion system; or a swirler is provided in the cooling duct.

25. The method according to claim 23, wherein a core duct passes through the catalyst chamber, at least in sections, through which duct oxygen flows only in the second operating mode of the rocket propulsion system.

26. A rocket propulsion system, comprising: a combustion chamber, an oxygen supply system configured to supply oxygen to the combustion chamber, a hydrogen supply system configured to supply hydrogen to the combustion chamber, an ignition system configured to initiate combustion of the oxygen-hydrogen mixture in the combustion chamber, and a control unit configured to control the oxygen supply system and the hydrogen supply system so that the rocket propulsion system is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber in a second mass mixing ratio of oxygen to hydrogen that is greater than the first mass mixing ratio.

27. The rocket propulsion system according to claim 26, wherein the first mass mixing ratio is a sub-stoichiometric mass mixing ratio of oxygen to hydrogen and wherein the second mass mixing ratio is a super-stoichiometric mass mixing ratio of oxygen to hydrogen.

28. The rocket propulsion system according to claim 27, wherein the first mass mixing ratio is less than or equal to 2.

29. The rocket propulsion system according to claim 27, wherein the second mass mixing ratio is greater than or equal to 50.

30. The rocket propulsion system according to claim 26, wherein the control unit is configured to control the oxygen supply system and the hydrogen supply system so that in the first operating mode of the rocket propulsion system, a first hydrogen mass flow is supplied to the combustion chamber that is greater than a second hydrogen mass flow that is supplied to the combustion chamber in the second operating mode of the rocket propulsion system, and wherein the control unit is configured to control the oxygen supply system and the hydrogen supply system so that in the second operating mode of the rocket propulsion system, the supply of hydrogen to the combustion chamber is interrupted.

31. The rocket propulsion system according to claim 26, wherein the control unit is configured to control the oxygen supply system and the hydrogen supply system so that in the first operating mode of the rocket propulsion system, a first oxygen mass flow is supplied to the combustion chamber that is smaller than a second oxygen mass flow that is supplied to the combustion chamber in the second operating mode of the rocket propulsion system.

32. The rocket propulsion system according to claim 26, wherein the control unit is configured to control the oxygen supply system, the hydrogen supply system and the ignition system so that at least a portion of the oxygen supplied to the combustion chamber and at least a portion of the hydrogen supplied to the combustion chamber are conducted into a catalyst chamber and the combustion of the oxygen-hydrogen mixture is initiated in the catalyst chamber, and at least one of: a flashback arrestor is arranged in the region of an entrance area to the catalyst chamber; the catalyst chamber comprises a premixing chamber for premixing of the oxygen conducted into the catalyst chamber and the hydrogen conducted into the catalyst chamber prior to initiating the combustion of the oxygen-hydrogen mixture; an oxygen supply opening for supplying oxygen to the catalyst chamber is formed in a wall of the catalyst chamber facing an oxygen supply duct; or a hydrogen supply duct for supplying hydrogen to the catalyst chamber opens into the catalyst chamber or a hydrogen supply opening for supplying hydrogen to the catalyst chamber is formed in a wall of the catalyst chamber facing a hydrogen supply duct.

33. The rocket propulsion system according to claim 32, wherein the catalyst chamber is enclosed, at least in sections, by a cooling duct arranged between an outer surface of the catalyst chamber and an inner surface of the combustion chamber, which duct opens into a combustion section of the combustion chamber arranged downstream of an exit area of the catalyst chamber and the control unit is configured to control the oxygen supply system or the hydrogen supply system so that oxygen supplied to the combustion chamber or hydrogen supplied to the combustion chamber flows through the cooling duct, and at least one of: the control unit is configured to control the oxygen supply system so that oxygen flows through the cooling duct only in the second operating mode of the rocket propulsion system; or a swirler is provided in the cooling duct.

34. The rocket propulsion system according to claim 32, wherein a core duct passes through the catalyst chamber, at least in sections, and the control unit is configured to control the oxygen supply system so that oxygen flows through the core duct only in the second operating mode of the rocket propulsion system.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

[0055] Preferred embodiments of the invention are explained in greater detail below with reference to the enclosed schematic drawings, wherein

[0056] FIG. 1 shows a schematic view of a rocket propulsion system,

[0057] FIG. 2 shows a diagram that illustrates a connection between the combustion temperature in a combustion chamber and a mixing ratio of an oxygen-hydrogen mixture to be ignited in the combustion chamber,

[0058] FIG. 3 shows an enlarged longitudinal section of a combustion chamber shown in FIG. 1 of the rocket propulsion system of a first embodiment,

[0059] FIG. 4 shows an enlarged longitudinal section of the combustion chamber shown in FIG. 1 of the rocket propulsion system of a second embodiment,

[0060] FIG. 5 shows an enlarged longitudinal section of the combustion chamber shown in FIG. 1 of the rocket propulsion system of a third embodiment,

[0061] FIG. 6 shows a schematic view of a rocket propulsion system of a fourth embodiment,

[0062] FIG. 7 shows an enlarged longitudinal section of a combustion chamber shown in FIG. 6 of the rocket propulsion system of the fourth embodiment,

[0063] FIG. 8 shows an enlarged longitudinal section of the combustion chamber shown in FIG. 8 of the rocket propulsion system of a fifth embodiment.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0064] FIG. 1 shows a rocket propulsion system 10, which comprises a combustion chamber 12, an oxygen supply system 14 connected to the combustion chamber 12 and a hydrogen supply system 16 connected to the combustion chamber. The oxygen supply system 14 is configured to supply oxygen to the combustion chamber 12. Correspondingly the hydrogen supply system 16 is configured to supply hydrogen to the combustion chamber 12. An ignition system 18 of the rocket propulsion system 10 is configured to initiate combustion of the oxygen-hydrogen mixture in the combustion chamber 12.

[0065] The rocket propulsion system 10 further comprises a control unit 20, which is configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that the rocket propulsion system 10 is operated alternately in a first operating mode, in which oxygen and hydrogen are supplied to the combustion chamber 12 in a first mass mixing ratio of oxygen to hydrogen, and in a second operating mode, in which oxygen and hydrogen are supplied to the combustion chamber 12 in a second mass mixing ratio of oxygen to hydrogen, which is greater than the first mass mixing ratio.

[0066] The oxygen supply system 14 comprises an oxygen supply line 22 connected to the combustion chamber 12, through which oxygen to be supplied to the combustion chamber 12 can flow. The oxygen supply line 22 is connected to an oxygen storage facility 24 of the oxygen supply system 14 for storing oxygen and is configured to supply oxygen from the oxygen storage facility 24 to the combustion chamber 12. An oxygen supply valve 26 connected to the control unit 20 is provided in the oxygen supply line 22, wherein an oxygen mass flow to be supplied to the combustion chamber 12 can be adjusted by means of the oxygen supply valve 26. The control unit 20 is configured to control the oxygen mass flow to be supplied to the combustion chamber 12 via the oxygen supply valve 26.

[0067] The hydrogen supply system 16 comprises a hydrogen supply line 28 connected to the combustion chamber 12, through which hydrogen to be supplied to the combustion chamber 12 can flow. The hydrogen supply line 28 is connected to a hydrogen storage facility 30 of the hydrogen supply system 16 for storing hydrogen and is configured to supply hydrogen from the hydrogen storage facility 30 to the combustion chamber 12. A hydrogen supply valve 32 connected to the control unit 20 is provided in the hydrogen supply line 28, by means of which a hydrogen mass flow to be supplied to the combustion chamber 12 via the hydrogen supply line 28 can be adjusted. The control unit 20 is configured to control the hydrogen mass flow to be supplied to the combustion chamber 12 via the hydrogen supply valve 32.

[0068] The oxygen supply system 14 and the hydrogen supply system 16 are connected to an electrolysis unit 34. The electrolysis unit 34 is configured to split water supplied to the electrolysis unit 34 from a water storage facility 36 via a water line 38 into hydrogen and oxygen by electrolysis. A water supply valve connected to the control unit 20 can further be provided in the water line, wherein a water mass flow to be supplied to the electrolysis unit 34 can be adjusted by means of the water supply valve. A non-return valve can also be arranged upstream of the water supply valve, i.e., opposite to the supply direction of the water. The oxygen produced in the electrolysis unit 34 can be supplied to the oxygen storage facility 24 via an oxygen line 40. The hydrogen produced in the electrolysis unit 34 can correspondingly be supplied via a hydrogen line 42 to the hydrogen storage facility 30. A non-return valve can also be provided in the oxygen line 40 and in the hydrogen line 42 respectively.

[0069] The combustion chamber 12 of the rocket propulsion system 10 is connected to a thruster 44, to which exhaust gases arising in the combustion chamber 12 due to combustion of the oxygen-hydrogen mixture can be supplied. The thruster 44 is provided to accelerate the exhaust gases produced in the combustion chamber 12 on their exit from the combustion chamber 12 up to an exit opening 46 of the thruster 44 and thereupon to discharge them to an environment of the rocket propulsion system 10 at high exit velocities, in order to generate thrust.

[0070] The control unit 20 is preferably configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that the combustion chamber 12 of the rocket propulsion system 10 is supplied in the first operating mode with oxygen and hydrogen in the first mass mixing ratio, wherein the first mass mixing ratio is a sub-stoichiometric mass mixing ratio of oxygen to hydrogen, in particular a mass mixing ratio of less than or equal to 2. In the second operating mode of the rocket propulsion system 10, the control unit 20 is preferably configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that the combustion chamber 12 of the rocket propulsion system 10 is supplied with oxygen and hydrogen in the second mass mixing ratio, wherein the second mass mixing ratio is a super-stoichiometric mass mixing ratio of oxygen to hydrogen, in particular a mass mixing ratio of greater than or equal to 50.

[0071] In particular, the control unit 20 can be configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that in the first operating mode of the rocket propulsion system 10, a first hydrogen mass flow is supplied to the combustion chamber 12, which is greater than a second hydrogen mass flow, which is supplied to the combustion chamber 12 in the second operating mode of the rocket propulsion system 10. In particular, the control unit 20 can be configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that the supply of hydrogen to the combustion chamber 12 is interrupted in the second operating mode of the rocket propulsion system 10.

[0072] Alternatively or in addition, the control unit 20 can be configured to control the oxygen supply system 14 and the hydrogen supply system 16 so that in the first operating mode of the rocket propulsion system 10, a first oxygen mass flow is supplied to the combustion chamber 12, which is smaller than a second oxygen mass flow, which is supplied to the combustion chamber 12 in the second operating mode of the rocket propulsion system 10.

[0073] FIG. 2 shows a diagram, which illustrates an exemplary operation of the rocket propulsion system 10 shown in FIG. 1 in the first operating mode and in the second operating mode of the rocket propulsion system 10. Here the combustion temperature in the combustion chamber 12 is shown as a function of the mass mixing ratio of oxygen to hydrogen of the oxygen-hydrogen mixture supplied to the combustion chamber 12. The abscissa of the diagram shows the mass mixing ratio of oxygen to hydrogen of the oxygen-hydrogen mixture supplied to the combustion chamber 12 and the ordinate of the diagram shows the combustion temperature in the combustion chamber 12 that arises due to combustion of the oxygen-hydrogen mixture supplied to the combustion chamber 12. As illustrated in FIG. 2, the combustion temperature in the combustion chamber 12 is at its highest during the combustion of a stoichiometric oxygen-hydrogen mixture, i.e., at a mass mixing ratio of oxygen to hydrogen of 8.

[0074] In the first operating mode of the rocket propulsion system 10, oxygen and hydrogen are supplied to the combustion chamber 12 in a mass mixing ratio of substantially 2. In the second operating mode of the rocket propulsion system 10, oxygen and hydrogen are supplied to the combustion chamber 12 in a mass mixing ratio of substantially 52. In the first and the second operating mode of the rocket propulsion system 10, the combustion temperature is 2000 K in each case.

[0075] FIG. 3 shows an enlarged longitudinal section of the combustion chamber 12 of the rocket propulsion system 10 shown in FIG. 1 with the ignition system 18 arranged therein. The ignition system 18 comprises a catalyst chamber 48 arranged in the combustion chamber 12, which catalyst chamber is configured to initiate the combustion of the oxygen-hydrogen mixture supplied to the combustion chamber 12. The control system 20 is configured to control the oxygen supply system 14, the hydrogen supply system 16 and the ignition system 18 so that at least a first portion of the oxygen supplied to the combustion chamber 12 and at least a portion of the hydrogen supplied to the combustion chamber 12 are conducted into the catalyst chamber 48 and the combustion of the oxygen-hydrogen mixture is initiated in the catalyst chamber 48. The catalyst chamber 48 has an entrance area 50, via which oxygen and hydrogen supplied to the combustion chamber 12 can be supplied to the catalyst chamber 48. In the region of the entrance area 50 of the catalyst chamber 48, a flashback arrestor 52 is arranged in the catalyst chamber 48, which arrestor is configured to prevent ignition upstream of the catalyst chamber 48 of the oxygen-hydrogen mixture to be supplied to the catalyst chamber 48.

[0076] Upstream of the flashback arrestor 52, the catalyst chamber 48 comprises a premixing chamber 54 for premixing the oxygen conducted into the catalyst chamber 48 and the hydrogen conducted into the catalyst chamber 48 prior to the initiation of combustion of the oxygen-hydrogen mixture. The premixing chamber 54 opens into the entrance area 50 to the catalyst chamber 48 and is arranged substantially perpendicular to this. A flow cross section of the premixing chamber 54 is formed so that it becomes larger in the direction of the entrance area 50 to the catalyst chamber 48, i.e., in the flow direction.

[0077] The catalyst chamber 48 further comprises an exit area 56, via which the exhaust gases produced in the catalyst chamber 48 are supplied to a combustion section 58 of the combustion chamber 12 arranged downstream of the exit area 56 of the catalyst chamber 48.

[0078] An oxygen supply duct 60 arranged in the combustion chamber 12 is configured to conduct at least the first portion of the oxygen supplied to the combustion chamber 12 into the premixing chamber 54 of the catalyst chamber 48 and a second portion of the oxygen supplied to the combustion chamber 12 into the combustion section 58 of the combustion chamber 12. To conduct the first portion of the oxygen supplied to the combustion chamber 12 into the premixing chamber 54 of the catalyst chamber 48, oxygen is supplied to the premixing chamber 54 of the catalyst chamber 48 via an oxygen supply opening 62, which is formed in a catalyst chamber wall 64 facing the oxygen supply duct 60.

[0079] Furthermore, a hydrogen supply duct 66 is arranged in the combustion chamber 12, via which the hydrogen supplied to the combustion chamber 12 is supplied to the catalyst chamber 48. The hydrogen supply duct 66 opens into the premixing chamber 54 of the catalyst chamber 48 and is arranged substantially perpendicular to the entrance area 50 of the catalyst chamber 48. In the embodiment of the rocket propulsion system 10 shown here, the hydrogen supply duct 66 has a circular cross section, wherein the hydrogen supply duct 66 passes through the oxygen supply duct 60. The oxygen supply duct 60 is formed accordingly in the form of an annular gap with an annular cross section.

[0080] The ignition system 18 of the rocket propulsion system 10 is formed so that the catalyst chamber wall 64 forms the catalyst chamber 48, the premixing chamber 54 and the hydrogen supply duct 66, wherein the oxygen supply duct 60 is arranged between an inner surface of the combustion chamber 12 and an outer surface of the catalyst chamber wall 64.

[0081] The catalyst chamber 48 is enclosed, at least in sections, by a cooling duct 68 arranged between an outer surface of the catalyst chamber 48 formed by the catalyst chamber wall 64 and the inner surface of the combustion chamber 12. The cooling duct 68 opens into the combustion section 58 of the combustion chamber 12 arranged downstream of the exit area 56 of the catalyst chamber 48 and the second portion of the oxygen supplied to the combustion chamber 12 flows through it. The oxygen supply duct 60 opens into the cooling duct 68.

[0082] Arranged in the cooling duct 68 is a swirl generation means in the form of a swirler 70, which is configured to induce swirl in the oxygen to be supplied to the combustion section 58 via the cooling duct 68. The dwell time in the cooling duct 68 of the oxygen flowing through the cooling duct 68 can thus be increased. Alternatively the swirl generation means can be provided in the form of tangentially positioned holes provided in the cooling duct 68.

[0083] A combustion chamber 12 of a second embodiment of the rocket propulsion system 10 is shown in FIG. 4. Compared with the embodiment shown in FIG. 3, an external outer surface of the catalyst chamber 48 adjoins the inner surface of the combustion chamber 12. Passing through the catalyst chamber 48, at least in sections, is a core duct 72, which opens into the combustion section 58 arranged downstream of the exit area 56 of the catalyst chamber 48. In the embodiment shown here, the oxygen supply duct 60 has a circular cross section, wherein the oxygen supply duct 60 passes through the hydrogen supply duct 66. The oxygen supply duct 60 opens into the core duct 72, wherein the second portion of the oxygen supplied to the combustion chamber 12 flows through the core duct 72 and is conducted by this into the combustion section 58 of the combustion chamber 12. The core duct 72 is configured to conduct the oxygen flowing through the core duct 72 along an internal outer surface of the catalyst chamber 48 formed by the catalyst chamber wall 64.

[0084] FIG. 5 shows an enlarged longitudinal section of the combustion chamber 12 shown in FIG. 1 of a rocket propulsion system 10 of a third embodiment. Here the core duct 72, through which the second portion of the oxygen supplied to the combustion chamber 12 flows and which conducts this into the combustion section 58 of the combustion chamber 12, passes through the catalyst chamber 48.

[0085] In the embodiment shown here, the hydrogen supply duct 66 is configured to conduct at least a first portion of the hydrogen supplied to the combustion chamber 12 into the catalyst chamber 48 and to conduct a second portion of the hydrogen supplied to the combustion chamber 12 via a cooling duct 74 into the combustion section 58 of the combustion chamber 12. The first portion of the hydrogen supplied to the combustion chamber 12 is supplied to the premixing chamber 54 of the catalyst chamber 48 via a hydrogen supply opening 76, which is formed in another, outer catalyst chamber wall 78 facing the hydrogen supply duct 66.

[0086] The cooling duct 74, through which the second portion of the hydrogen supplied to the combustion chamber 12 can flow, is arranged between the external outer surface of the catalyst chamber 48 and the inner surface of the combustion chamber 12. The catalyst chamber 48 is enclosed, at least in sections, by the cooling duct 74. The cooling duct 74 is formed so that the hydrogen supplied to the combustion section 58 via the cooling duct 74 forms a reactive and low-oxidizer cooling film on the inner surface of the combustion chamber 12 in the combustion section 58 of the combustion chamber 12. This has the effect that reductive conditions can prevail along an inner surface of the combustion chamber 12 and thus a reaction of a combustion chamber wall with oxygen can be prevented. At the same time, lower combustion temperatures can be reached by this in the region of the combustion chamber wall in operation of the rocket propulsion system 10. Furthermore, the core duct 72 is formed so that the oxygen supplied to the combustion section 58 of the combustion chamber 12 via the core duct 72 forms an oxidizer-rich gas core with higher combustion temperatures in the combustion section 58 of the combustion chamber 12, which core is enclosed by the cooling film formed by the cooling duct 74. In other words, due to the cooling film enclosing the gas core, a thermal insulation can be provided between the gas core having high combustion temperatures and the combustion chamber wall. The thermal loading on the combustion chamber wall during the operation of the rocket propulsion system 10 can thus be reduced.

[0087] A swirl generation means in the form of a swirler 80 is arranged in the cooling duct 74 and is configured to induce swirl in the hydrogen to be supplied to the combustion section 58 via the cooling duct 74. Alternatively the swirl generation means can be provided in the form of tangentially positioned holes provided in the cooling duct 74.

[0088] A schematic view of a rocket propulsion system 10 of a fourth embodiment is shown in FIG. 6. The rocket propulsion system 10 shown in FIG. 6 differs from the embodiment shown in FIG. 1 in that the oxygen supply system 14 has a further oxygen supply line 82 connected to the combustion chamber 12. The further oxygen supply line 82 is configured to supply oxygen from the oxygen storage facility 24 to the combustion chamber 12. Provided in the further oxygen supply line 82 is a further oxygen supply valve 84 connected to the control unit 20, wherein a further oxygen mass flow to be supplied to the combustion chamber 12 can be adjusted via the further oxygen supply valve 84.

[0089] FIG. 7 shows an enlarged longitudinal section of the combustion chamber 12 shown in FIG. 6. Compared with the embodiment of the rocket propulsion system 10 shown in FIG. 3, the oxygen supply duct 60 arranged in the combustion chamber 12 is not connected to the cooling duct 68 in the embodiment shown here. In other words, the oxygen supply duct 60 is separate from the cooling duct 68. The oxygen supply duct 60 is configured to supply the oxygen, which is supplied to the combustion chamber 12 by the oxygen supply line 22 of the oxygen supply system 14, to the catalyst chamber 48 via the oxygen supply opening 62 and the premixing chamber 54 of the catalyst chamber 48.

[0090] In the embodiment shown here, the cooling duct 68 is connected to the further oxygen supply line 82 of the oxygen supply system 14 and is configured to supply the oxygen, which is supplied to the combustion chamber 12 by the further oxygen supply line 82 of the oxygen supply system 14, to the combustion section 58 of the combustion chamber 12.

[0091] The control unit 20 is preferably configured in this case to control the oxygen supply system 14, the hydrogen supply system 16 and the ignition system 18 so that the oxygen mass flow flowing through the cooling duct 68 is varied in the first operating mode and the second operating mode of the rocket propulsion system 10. In particular, the control unit 20 can be configured to control the oxygen supply system 14, the hydrogen supply system 16 and the ignition system 18 so that oxygen flows through the cooling duct 68 only in the second operating mode of the rocket propulsion system 10.

[0092] FIG. 8 shows an enlarged longitudinal section of the combustion chamber 12 shown in FIG. 6 of the rocket propulsion system 10 of a fifth embodiment. Compared with the embodiment of the rocket propulsion system 10 shown in FIG. 4, the oxygen supply duct 60 arranged in the combustion chamber 12 is not connected to the core duct 72. In other words, the oxygen supply duct 60 is provided so that it is separate from the core duct 72. The oxygen supply duct 60 is configured to supply the oxygen, which is supplied to the combustion chamber 12 by the oxygen supply line 22 of the oxygen supply system 14, to the catalyst chamber 48 via the oxygen supply opening 62 and the premixing chamber 54 of the catalyst chamber 48.

[0093] In the embodiment shown here, the core duct 72 is connected to the further oxygen supply line 82 of the oxygen supply system 14 and is configured to supply the oxygen, which is supplied to the combustion chamber 12 by the further oxygen supply line 82, to the combustion section 58 of the combustion chamber 12.

[0094] The control unit 20 is preferably configured in this case to control the oxygen supply system 14, the hydrogen supply system 16 and the ignition system 18 so that the oxygen mass flow flowing through the core duct 72 is varied in the first operating mode and the second operating mode of the rocket propulsion system 10. In particular, the control unit 20 can be configured to control the oxygen supply system 14, the hydrogen supply system 16 and the ignition system 18 so that oxygen flows through the core duct 72 only in the second operating mode of the rocket propulsion system 10.

[0095] While at least one exemplary embodiment of the present invention(s) is disclosed herein, it should be understood that modifications, substitutions and alternatives may be apparent to one of ordinary skill in the art and can be made without departing from the scope of this disclosure. This disclosure is intended to cover any adaptations or variations of the exemplary embodiment(s). In addition, in this disclosure, the terms “comprise” or “comprising” do not exclude other elements or steps, the terms “a” or “one” do not exclude a plural number, and the term “or” means either or both. Furthermore, characteristics or steps which have been described may also be used in combination with other characteristics or steps and in any order unless the disclosure or context suggests otherwise. This disclosure hereby incorporates by reference the complete disclosure of any patent or application from which it claims benefit or priority.