Tumbler Device and Method for Operating a Tumbler Device

Description

[0001] The present invention relates to a thrust chamber device comprising a thrust chamber with a thrust space, which has a first section, a second section adjoining the second section, and a third section adjoining the second section, wherein the thrust space in all three sections is bounded by an outer nozzle wall with an outer thrust space surface, which outer thrust space surface tapers towards the third section in the first and second sections and widens away from the second section in the third section, wherein a narrowest point is formed at the transition from the second section to the third section, wherein the first section is bounded by an inner nozzle wall with an inner thrust space surface which tapers towards the second section, and wherein an annular combustion chamber is formed between the inner thrust space surface and the outer thrust space surface, which extends over the first section.

[0002] Furthermore, the present invention relates to a method for operating a thrust chamber device.

[0003] Furthermore, the invention relates to an engine, in particular for a missile or aircraft.

[0004] Furthermore, the present invention relates to a missile or aircraft comprising a first fuel storage device for at least a first fuel component, a second fuel storage device for at least a second fuel component and an engine.

[0005] Thrust chamber devices of the type described above are used, for example, in jet engines to generate thrust, such as to propel a missile like a rocket, particularly through the combustion of propellant components. A propellant component can be, in particular, a fuel. For example, liquid hydrogen (LH₂) can be used as the fuel, and liquid oxygen (LOX), which acts as an oxidizer, can be used as another propellant component.

[0006] An example of a thrust chamber device of the type described above is known in particular from WO 2018 / 167204 A1. In the known thrust chamber device, a plurality of first fuel inlets are formed in the region of the outer nozzle wall to introduce a first fuel component through them into the thrust chamber. The outer nozzle wall is simultaneously cooled by the first fuel component, which heats up in the process. However, with such transpiration cooling, it is necessary to inject more fuel into the thrust chamber than is required for cooling the outer nozzle wall. As a result, the first fuel component is not completely combusted in the thrust chamber or combustion chamber.

[0007] WO 02 / 29233 A1 discloses a ring-shaped combustion chamber for a rocket engine powered by a liquid propellant. GB 2 196 394 A describes a rocket engine.

[0008] It is therefore an object of the present invention to improve a thrust chamber device, a method, an engine and a missile or aircraft of the type described above in such a way that the thrust chamber device can be operated more efficiently.

[0009] This problem is solved according to the invention in a thrust chamber device of the type described above by the fact that the thrust chamber device comprises a regenerative cooling device for cooling the inner nozzle wall and the outer nozzle wall with a coolant, in order to fully utilize the heat or enthalpy absorbed by the coolant during the cooling of the two nozzle walls for driving a drive device with which conveying devices for fuel components to be burned in the thrust chamber can be driven.

[0010] Regenerative cooling of the inner and outer nozzle walls with the coolant offers the particular advantage that the heat or enthalpy absorbed during cooling of the two nozzle walls can be fully utilized, for example, to drive a turbine that powers conveying devices, especially turbopumps, for the fuel components to be combusted in the combustion chamber. Furthermore, the proposed regenerative cooling eliminates the need to inject more of the first fuel component into the combustion chamber than is required for optimal combustion. Additional advantages arise for a regenerative cooling system in conjunction with the specially shaped combustion chamber. In particular, the length of the combustion chamber can be significantly reduced compared to a conventional combustion chamber design with a cylindrical combustion chamber wall.Optimization in this respect can be achieved particularly through the shape of the inner and outer nozzle walls, for example, by shaping both nozzle walls as hyperboloids of revolution. Length reductions, and thus significant weight savings of up to 50% for the thrust chamber assembly, offer considerable advantages, especially in the aerospace sector. In particular, this allows for the transport of a larger payload with the same thrust chamber performance. The shorter design, i.e., a reduced overall length compared to conventional combustion chamber designs, also prevents excessive pressure loss due to enthalpy absorption by the coolant as it flows through the nozzle wall. Furthermore, it is also possible, for example, to divide the coolant for cooling between the inner and outer nozzle walls, thus cooling both walls in parallel, or to perform this cooling of the nozzle walls serially.This allows for the use of different cooling modes, optimally adapted to the different operating modes of the thrust chamber device. In particular, serial cooling of the inner and outer nozzle walls is not possible at all with transpiration cooling.

[0011] It is advantageous if the regenerative cooling system comprises a plurality of inner coolant channels in the inner nozzle wall and a plurality of outer coolant channels in the outer nozzle wall, and if the plurality of inner and outer coolant channels are designed to allow coolant flow through them. This enables efficient cooling of the nozzle walls. The coolant channels can extend essentially parallel to a longitudinal or symmetry axis of the thrust chamber, or they can, for example, be spirally shaped in the nozzle walls relative to the longitudinal or symmetry axis. Heat absorption in the coolant, and in particular pressure loss due to heat absorption during cooling, can thus be optimized depending on the size of the thrust chamber device.

[0012] Cooling channels can be formed in the nozzle walls of the thrust chamber device, particularly in a simple manner, if the inner nozzle wall and / or the outer nozzle wall are made of a ceramic and / or metallic material.

[0013] Preferably, the inner shear chamber surface and / or the outer shear chamber surface have at least partially the shape of a hyperboloid of revolution or a continuously concave longitudinal section. In particular, the inner shear chamber surface and / or the outer shear chamber surface can have a shape similar to that of a hyperboloid of revolution, insofar as they exhibit a continuously concave longitudinal section. For example, regeneratively cooled double-shell hyperboloid shear chambers or combustion chambers can be designed in this way. In particular, it is possible in a simple manner to define a constant cross-section of the shear chamber; for example, the shear chamber can be designed as an annular space between the outer shear chamber surface and the inner shear chamber surface in the first section. Furthermore, it is possible to delimit the different sections of the shear chamber by shear chamber surfaces that define differently curved hyperboloids of revolution.

[0014] To achieve efficient cooling and guidance of the combustion products in the thrust chamber, it is advantageous for the inner thrust chamber surface to be convexly curved or substantially convexly curved in the direction of the outer thrust chamber surface. Substantially convexly curved means, in particular, that an envelope of the inner thrust chamber surface is convexly curved. The inner thrust chamber surface can then also have short areas of concave curvature, for example, to modify the cross-sectional area of ​​the annular combustion chamber, especially to increase it in order to form fuel mixing zones.

[0015] It is advantageous if the outer thrust chamber surface is convexly curved or substantially convexly curved in the direction of the inner thrust chamber surface. Compared to a classic contour of an outer nozzle wall of a thrust chamber with a substantially cylindrical combustion chamber, this avoids typical boundary layer turbulence, also known as Görtler vortices. Such vortices form particularly in the region of the narrowest point of the thrust chamber, which is defined by the outer nozzle wall.

[0016] According to a further preferred embodiment of the invention, the thrust chamber can be configured to define a longitudinal axis and that the thrust chamber, in particular the first section and / or the second section and / or the third section, is rotationally symmetrical with respect to the longitudinal axis. This particularly simplifies the design and construction of the thrust chamber device.

[0017] It is advantageous if the outer shear chamber surface and / or the inner shear chamber surface are rotationally symmetrical with respect to the longitudinal axis. This can, for example, simplify the design of the shear chamber device.

[0018] Preferably, the annular combustion chamber has a constant or substantially constant cross-sectional area. Although the diameters of the inner and outer thrust chamber surfaces can vary along the thrust chamber's extension parallel to the longitudinal axis, a quasi-cylindrical annular combustion chamber can be formed in this way. This convergent thrust chamber section for sound transmission transitions into a convergent section without changing the curvature direction of the thrust chamber surfaces, due to the geometric interaction of the purely convex outer and purely convex inner thrust chamber surfaces. In particular, the tapered outer nozzle wall can prevent excessive heating of the nozzle, especially by increasing the coolant flow, for example, by enlarging the coolant channels in the region of the smallest diameter of the outer thrust chamber surface.The constant cross-sectional area makes it possible in particular to form a quasi-cylindrical combustion chamber zone without cross-sectional narrowing, i.e. a combustion chamber zone analogous to the standard design, namely the cylindrical part of a rocket combustion chamber.

[0019] To enable a controlled, and in particular stoichiometric, chemical reaction of the propellant components in the thrust chamber, it is advantageous for the thrust chamber device to include a plurality of first propellant inlets for a first propellant component and a plurality of second propellant inlets for a second propellant component. A plurality of propellant inlets for the different propellant components allows, in particular, optimal mixing of these components upon or near their entry into the thrust chamber.

[0020] It is advantageous if the first section of the thrust chamber is bounded at an end pointing away from the second section by an injection wall that connects the inner and outer nozzle walls, and if the majority of first and second fuel inlets are arranged or formed in the injection wall. In particular, all fuel inlets in the thrust chamber assembly can be arranged or formed in the injection wall. This makes it possible, in particular, to introduce the fuels into the thrust chamber in a direction predetermined by the annular combustion chamber, towards the third section.

[0021] Preferably, the injection wall is ring-shaped, rotationally symmetrical, or hyperboloid-like for sealing the ring-shaped combustion chamber. This makes it particularly easy to seal the thrust chamber proximally, i.e., at its end furthest from the third section.

[0022] To inject the fuel components into the combustion chamber in a defined manner, it is advantageous for the combustion chamber device to include an injection head, and for the injection head to include the injection wall. In this way, the injection head and the injection wall can, in particular, seal the annular combustion chamber. The injection head can be designed with one, two, or more distribution chambers to introduce the different fuel components into the combustion chamber in a targeted manner and simultaneously through the multiple fuel inlets.

[0023] It is advantageous if the majority of the first fuel inlets and the majority of the second fuel inlets are designed in the form of channels with channel openings pointing into the annular combustion chamber. This allows the fuel components to be introduced into the annular combustion chamber in a defined and simple manner, particularly with a flow component already running in a direction parallel or substantially parallel to the thrust chamber surfaces bounding the annular combustion chamber.

[0024] It is advantageous if the majority of the first fuel inlets define the first fuel inlet longitudinal axes and if these axes point into the first section in a direction parallel or substantially parallel to tangents to the inner thrust chamber surface and / or the outer thrust chamber surface. The first fuel component can thus be introduced into the annular combustion chamber in the direction defined by the fuel inlet longitudinal axes and flow through the thrust chamber.

[0025] Preferably, the thrust chamber assembly comprises a first injection device for injecting the at least one first propellant component into the thrust chamber through the plurality of first propellant inlets. The first injection device allows the at least one first propellant component to be introduced into the thrust chamber in a defined manner, in particular with a predefined volumetric flow rate.

[0026] It is advantageous if the thrust chamber device includes a first fuel storage unit for a first fuel component and if the first injection device includes a first delivery device for conveying the at least one first fuel component from the first fuel storage unit through the plurality of first fuel inlets into the thrust chamber. In this way, the first fuel component can, for example, be injected directly from the first fuel storage unit into the thrust chamber through the first fuel inlets, or, prior to being introduced into the thrust chamber, the inner and / or outer nozzle wall can be cooled if the first fuel component is guided, for example, through inner and / or outer coolant channels in the nozzle walls before being introduced into the thrust chamber. The first delivery device can, in particular, be designed in the form of a pumping device, for example, a turbopump.

[0027] It is advantageous if the first conveying device has a suction side and a pressure side, if the suction side is fluidly connected to the first fuel storage tank, and if the pressure side is fluidly connected to the outer coolant channels and / or the inner coolant channels. This design makes it possible, in particular, to convey the first fuel component through the inner and / or outer coolant channels using the first conveying device.

[0028] According to a further preferred embodiment of the invention, the plurality of second fuel inlets define second fuel inlet longitudinal axes, and these axes extend into the first section in a direction parallel or substantially parallel to tangents to the inner thrust chamber surface and / or the outer thrust chamber surface. This allows the at least one second fuel component to be injected into the thrust chamber with high efficiency and low friction losses. In particular, this makes it possible to introduce the first fuel component and the at least one second fuel component into the thrust chamber parallel to each other, already with an optimized flow direction predetermined by the thrust chamber.

[0029] It is advantageous if the thrust chamber device includes a second injection device for injecting the at least one second fuel component into the thrust chamber through the plurality of second fuel inlets. With the second injection device, the at least one second fuel component can be introduced into the thrust chamber in a defined manner, for example, with a predefined volume flow rate. In particular, the second injection device and the first injection device can be operated in a coordinated manner to ensure that an optimal fuel mixture, especially a stoichiometric mixture, is always maintained in the thrust chamber for combustion.

[0030] It is advantageous if the thrust chamber device includes at least one second fuel storage unit for at least one second fuel component, and if the second injection device includes a second delivery device for conveying the at least one second fuel component from the second fuel storage unit through the plurality of second fuel inlets into the thrust chamber. In this way, the second fuel component can be introduced into the thrust chamber simply and in a defined manner, for example, with a predetermined volume flow rate. The second fuel component can, in particular, be a liquid oxidizer. The second delivery device can, in particular, be designed in the form of a pump, for example, a turbopump.

[0031] Advantageously, the first propellant component also acts as the coolant. This allows for simple and precise cooling of the thrust chamber device. In particular, this avoids the need for a complex nozzle wall design, as, for example, coolant channels only need to be provided for one of the propellant components to operate the thrust chamber device. This effectively and reliably prevents any undesirable reaction between the propellant components before they enter the thrust chamber.

[0032] Optimal cooling of the nozzle walls can be achieved particularly when the first fuel component is a liquid fuel. This can be, in particular, liquid hydrogen, liquid methane, or liquefied natural gas. As the first fuel component flows through the nozzle walls, its pressure increases due to heat absorption. This pressure increase can, in particular, drive a drive unit for the conveying equipment, such as a turbine that is effectively coupled to the conveying equipment.

[0033] Advantageously, the second fuel component is a liquid oxidizer. For example, this could be liquid oxygen. Large volumes of gas, especially in liquid form, can be stored compactly, for example in a pressure tank.

[0034] It is advantageous if the thrust chamber device includes a drive unit for powering the first conveying unit and / or the second conveying unit. In particular, two drive units can be provided, each assigned to one of the two conveying units. Specifically, one drive unit can be coupled to both conveying units, for example, via a common drive shaft. The described configurations make it possible, in particular, to use one of the two fuel components not only for combustion and thus for generating thrust with the thrust chamber device, but also to convey the fuel components directly or indirectly from the respective fuel storage unit into the thrust chamber, i.e., in particular via coolant channels.

[0035] The drive unit can easily be designed in the form of a turbine. For example, it can be designed in the form of a gas turbine.

[0036] A simple and compact design of the thrust chamber device can be achieved in particular by the fact that the drive unit includes a fluid inlet and a fluid outlet and that the fluid outlet is fluidly connected to the first fuel inlets.

[0037] According to a further preferred embodiment of the invention, the outer coolant channels may have outer coolant channel inlets and outer coolant channel outlets, and the inner coolant channels may have inner coolant channel inlets and inner coolant channel outlets. This configuration makes it particularly possible to equip the nozzle walls with coolant channels as desired. The respective coolant channel inlets and coolant channel outlets may optionally be arranged or configured either near an end of the first section pointing away from the third section or at a distance from this end.For example, coolant can be directed through external coolant channels from the third section to the first section, and from the end of the first section facing away from the third section, through internal coolant channels to the end of the first section facing the second section. Alternatively, a coolant flow can also be specified in the reverse direction.

[0038] It is advantageous if the pressure side of the first conveying device is fluidly connected to the outer coolant channel inlets, if the outer coolant channel outlets are fluidly connected to the inner coolant channel inlets, and if the inner coolant channel outlets are fluidly connected to the fluid inlet of the drive unit. Such a configuration particularly enables serial cooling of the outer and inner nozzle walls. This means, in particular, that the coolant first flows through the outer nozzle wall, then the inner nozzle wall, and then is directed to the drive unit. Providing a coolant flow in this manner is particularly advantageous when starting or starting the thrust chamber device, since the coolant heats up more effectively overall due to the longer flow path and can thus drive the drive unit with high efficiency.

[0039] Furthermore, it can be advantageous if the pressure side of the first conveying device is fluidly connected to the outer and inner coolant channel inlets, and if the outer and inner coolant channel outlets are fluidly connected to the fluid inlet of the drive unit. Such a configuration allows, in particular, a quasi-parallel coolant flow through both the outer and inner coolant channels. The coolant absorbs less heat in this configuration than with the serial cooling of the two nozzle walls described above. However, with parallel cooling of the nozzle walls, friction is reduced due to the shorter flow path of the coolant through each nozzle wall.Parallel cooling of the nozzle walls is particularly advantageous after the start-up of the thrust chamber device, i.e., especially during normal or load operation of the thrust chamber device.

[0040] Furthermore, it can be advantageous if the pressure side of the first conveying device is fluidly connected to the inner coolant channel inlets, if the inner coolant channel outlets are fluidly connected to the outer coolant channel inlets, and if the outer coolant channel outlets are fluidly connected to the fluid inlet of the drive device. In this way, serial cooling of the nozzle walls can be achieved, in particular by first cooling the inner nozzle wall with coolant from the first fuel reservoir, and then using the preheated coolant to cool the outer nozzle wall, thereby heating it further. Thus, serial cooling of the nozzle walls can also be achieved in this way.

[0041] It is advantageous if the thrust chamber device includes a coolant flow switching device for selectively switching the coolant flow between parallel and series through the inner and outer coolant channels. Such a coolant flow switching device makes it possible, in particular, to selectively cool the nozzle walls in parallel or series, as explained in detail above. Specifically, the coolant flow switching device can be designed such that it can also be specified whether the series cooling should first occur to the inner or outer nozzle wall.

[0042] The coolant flow switching device can be easily designed if it includes a valve device.

[0043] The problem set out at the beginning is further solved in an engine of the type described at the beginning according to the invention by comprising one of the thrust chamber devices described above.

[0044] Such a further developed engine exhibits the advantages already described above in connection with preferred embodiments of thrust chamber devices.

[0045] The aforementioned problem is further solved according to the invention in a missile or aircraft of the type described above by comprising an engine according to the invention.

[0046] The improved cooling of the nozzle walls of the thrust chamber assembly, combined with the compact design of the thrust chamber assembly, allows for the development of shorter and therefore significantly lighter engines with the same system performance. This makes it possible to move larger payloads with the same amount of fuel.

[0047] The problem set out at the beginning is further solved in a method of the type described at the beginning according to the invention by regeneratively cooling the inner nozzle wall and the outer nozzle wall with a coolant in order to fully utilize the heat or enthalpy absorbed by the coolant during the cooling of the two nozzle walls for driving a drive device with which one or more conveying devices for fuel components to be burned in the thrust chamber can be driven.

[0048] As described in detail above, this method allows for optimal cooling of the thrust chamber device without injecting an excess of the first propellant component into the thrust chamber. This minimizes mass flow losses of the first propellant component and results in more efficient operation of the thrust chamber device overall.

[0049] Advantageously, the first fuel component is used as a coolant. This eliminates the need to carry a separate coolant solely for cooling purposes. The first fuel component can then cool the nozzle walls before reacting with the next fuel component, thus fulfilling a dual function.

[0050] It is advantageous if the coolant is guided in parallel through the outer and inner nozzle walls. This is particularly beneficial during normal or load operation of the thrust chamber device as described above. Guiding in parallel through the nozzle walls means, in particular, splitting the coolant flow so that part of it is guided through the outer nozzle wall and another part through the inner nozzle wall. This approach is especially advantageous when the engine is in a stable operating mode, i.e., particularly during normal or load operation. This minimizes friction losses as the coolant flows through the nozzle walls.

[0051] Alternatively, it can be advantageous to guide the coolant first through the inner nozzle wall and then through the outer nozzle wall, or first through the outer nozzle wall and then through the inner nozzle wall. This achieves a so-called serial regenerative cooling of the nozzle walls. In this procedure, the coolant flow is initially guided through one of the two nozzle walls and then through the other. Such a cooling mode is particularly advantageous when starting or initiating the thrust chamber device, as the coolant can heat up more due to the longer flow path through the nozzle walls. The energy absorbed in the coolant can then be used, in particular, to power a drive unit for fuel delivery systems.This described cooling mode is particularly advantageous when starting or starting the thrust chamber device.

[0052] Preferably, the cooling mode is changed when the operating mode of the thrust chamber device is changed. As already described, serial cooling of the nozzle walls can be used, for example, when the thrust chamber device is started. When a load mode of the thrust chamber device is reached, the system can then switch to parallel cooling of the nozzle walls, as described.

[0053] It is advantageous to use a serial cooling mode in the start-up mode of the thrust chamber device and a parallel cooling mode in the load mode. Parallel and serial cooling have already been described in detail above. They define the flow through the nozzle walls either sequentially (serially) or in parallel (partially through the inner nozzle wall and partially through the outer nozzle wall). This approach allows for optimized operation of the thrust chamber device, as the cooling of the nozzle walls can be optimally tailored to the respective operating mode.

[0054] It is advantageous if the coolant flows through the drive unit before or after passing through the inner and outer nozzle walls. The enthalpy absorption of the coolant by absorbing heat from the nozzle walls can then be used, in particular, to drive the drive unit, which is intended to power one or more conveying devices for the required fuel components.

[0055] Preferably, after flowing through the inner and / or outer nozzle wall, the coolant is injected into the thrust chamber via the first fuel inlets. This allows the coolant, and in particular the first fuel component, to be used not only to cool the thrust chamber but also to generate thrust.

[0056] The following description of preferred embodiments of the invention, in conjunction with the drawings, serves for further explanation. The drawings show: Figure 1: A schematic representation of a thrust chamber contour with two hyperboloids defining the inner and outer nozzle walls; Figure 2: A schematic representation of an embodiment of a thrust chamber device with serial regenerative cooling of the outer and inner nozzle walls; Figure 3: A schematic representation of another embodiment of a thrust chamber device with parallel regenerative cooling of the inner and outer nozzle walls; Figure 4: A schematic representation of another embodiment of a thrust chamber device with parallel cooling of the inner and outer nozzle walls; Figure 5: A schematic representation of another embodiment of a thrust chamber device with serial cooling of the inner and outer nozzle walls; Figure 6: A schematic representation of an embodiment of an injection head of a thrust chamber device;Figure 7: a schematic representation of a coolant flow switching device; Figure 8: a schematic representation of an engine; and Figure 9: a schematic representation of a missile or aircraft.

[0057] A first embodiment of a thrust chamber device 10 is shown schematically in Figure 1 The thrust chamber device 10 comprises a thrust chamber 12 with a first section 14, a second section 16 and a third section 18, which divide a thrust space 20 of the thrust chamber 10 into corresponding sections 14, 16 and 18.

[0058] The first section 14 is defined by an insert body 22 placed in the thrust chamber 20, which defines an inner nozzle wall 24 with an inner thrust chamber surface 26 of the thrust chamber 20.

[0059] The thrust chamber 20 is bounded on the outside by an outer nozzle wall 28, which is rotationally symmetric with respect to a longitudinal axis 30, in the form of a single-shell hyperboloid of revolution.

[0060] The inner shear space surface 26 has the shape of a two-shell hyperboloid of revolution. A tangent 32 at the transition between the first section 14 and the second section 16 to the insertion body 22 runs perpendicular to the longitudinal axis 30.

[0061] The first section 32 defines an annular combustion chamber 34 as part of the thrust chamber 20.

[0062] The outer nozzle wall 28 defines an outer thrust chamber surface 36 that limits the thrust chamber 20.

[0063] In one embodiment, the shear space areas 26 and 36 are selected such that a free cross-sectional area 38 is constant depending on a distance 40 from a first end 42 of the first section 14. The first end 42 limits the first section 14 away from the second section 16.

[0064] Starting from the first end 42, the outer diameter of the ring surface defined by the annular combustion chamber 34 decreases towards the second section 16, as does the diameter of an inner boundary of this ring surface formed by the inner thrust chamber surface 26. In the transition 54 between the first section 14 and the second section 16, the ring surface then transitions without curvature into a circular surface, the cross-section of which continues to taper towards the narrowest point 46 of the thrust chamber 20.

[0065] The narrowest point 46 is defined by a minimum diameter of the outer nozzle wall 28. Through the narrowest point 46, combustion gases are accelerated divergently towards a nozzle outlet 48. The third section 18, which extends from the narrowest point 46 to the nozzle outlet 48, is also referred to as the supersonic region.

[0066] The continuous curvature of the outer nozzle wall 28 virtually eliminates the disadvantages of a classic thrust chamber profile with a cylindrical combustion chamber and a subsequent Lavalle nozzle towards the nozzle outlet. With the classic profile, boundary layer turbulence typically occurs in the area of ​​the outer nozzle wall where its curvature direction changes. In contrast, with the profile described in Figure 1 In the illustrated embodiment, the outer nozzle wall 28 is continuously curved, specifically convex in the direction of the longitudinal axis 30.

[0067] The inner nozzle wall 24 has a continuously convexly curved inner thrust chamber surface 26 pointing away from the insertion body 22.

[0068] Through the as described at the in Figure 1 In the thrust chamber 12 shown, with a constant cross-section of the annular combustion chamber 34, a quasi-cylindrical annular combustion chamber 34 is formed, but with the advantage that the outer thrust chamber surface 36 tapers continuously from the first end 42, so that neither separation nor turbulence can occur in the area of ​​the narrowest point 46.

[0069] By appropriately selecting a shape for the shear space surfaces 26 and 36, the cross-sectional area 38 can be constant as described or alternatively increase or decrease in the direction of the transition 44.

[0070] The second section 16 extends from the transition 44 to the narrowest point 46 and is limited exclusively by the outer thrust space area 36.

[0071] The first section 14 forms the basis for the in Figure 1 In the illustrated embodiment, a quasi-cylindrical region is formed. The second section 16 forms a convergent region of the thrust chamber 20, and the third section 18 forms the divergent supersonic region of the thrust chamber 12.

[0072] Another embodiment of a thrust chamber device designated in its entirety by reference numeral 10 is shown schematically in Figure 2 depicted. Both in Figure 2 as well as in the further Figures 3 to 9 are for easier understanding of identical or similar components of the thrust chamber device 10, which are used in conjunction with Figure 1 The same reference symbols were used in the description.

[0073] The thrust chamber device 10 comprises a regenerative cooling device 50 for cooling the inner nozzle wall 24 and the outer nozzle wall 28 with a coolant 52.

[0074] The thrust chamber device 10 comprises a first fuel storage container 54 for a first fuel component 56 and a second fuel storage container 58 for a second fuel component 60.

[0075] In the described embodiment, a liquid fuel 62 is used as the first fuel component 56. An oxidizer 64, or an oxidizing agent, is used as the second fuel component 60. Both fuel components 56 and 60 are stored in liquid form and at cryogenic temperatures in the fuel storage tanks 54 and 58, respectively.

[0076] At the in Figure 2 In the schematically illustrated embodiment of the thrust chamber device 10, the first fuel component 56 is the coolant 52.

[0077] The first fuel storage tank 54 is fluidly connected via a connecting line 66 to a first pumping device 68, specifically to a suction side 70 thereof. A pressure side 72 of the first pumping device 68 is fluidly connected via a further connecting line 74 to a ring distributor 76, which annularly surrounds the outer nozzle wall 28 in the region of the third section 18.

[0078] From the ring distributor 76, a plurality of outer coolant channels 78 extend in the outer nozzle wall 28, which are fluidly connected to the ring distributor 76 via their outer coolant channel inlets 80, to the first end 42 of the thrust chamber device 10.

[0079] The outer coolant channel outlets 82 of the outer coolant channels 78 are fluidly connected via connecting lines 84 to the inner coolant channel inlets 86 of the inner coolant channels 88.

[0080] The inner coolant channels 88, of which a plurality are provided, extend in the inner nozzle wall 24 to near a distal end 90 of the insert body 22, i.e., to near the second section 16. Inner coolant channel outlets 94 of the inner coolant channels 88 are fluidly connected to one another via a collector 92.

[0081] The collector 92 is fluidly connected via a further connecting line 96 to a fluid inlet 98 of a drive unit 100.

[0082] A fluid outlet 102 of the drive unit 100, which is designed in the form of a turbine 104, is fluidly connected to a distribution chamber 108 via a connecting line 106. From the distribution chamber 108, a plurality of first fuel inlets 110 extend through an injection wall 112 into the first section 114 of the thrust chamber 20.

[0083] In the manner described, the first fuel component 56 can be conveyed from the first fuel storage tank 54 by means of the first conveying device 68 through the coolant channels 78 and 80 to the drive unit 100. As the first fuel component 56 flows through the outer nozzle wall 28 and the inner nozzle wall 24, it is heated, expands, and drives the turbine 104, which is effectively connected to the first conveying device 68 and a second conveying device 114. This is shown schematically in Figure 2 in the form of a drive shaft 116.

[0084] The first fuel component 56 thus serves not only as fuel but also as coolant 52. In the described manner, a serial regenerative cooling of the nozzle walls 28 and 24 is realized, namely first cooling of the outer nozzle wall 28 and subsequently cooling of the inner nozzle wall 24 with the coolant 52, which is then already somewhat warmed.

[0085] The serial cooling mode of the cooling device 50 described above is particularly advantageous when starting the thrust chamber device 10. The coolant 52 is heated over a longer path and can thus absorb a greater amount of heat, which is advantageous for operating the drive device 104.

[0086] The second fuel storage tank 58 is fluidly connected to the second pumping device 114 via a connecting line 118, specifically to its suction side 120. A pressure side 124 of the second pumping device 114 is fluidly connected to a distribution chamber 126 of an injection head 128 via a further connecting line 122. The distribution chamber 126 for the second fuel component 60 is fluidly connected to the first section 14 of the thrust chamber 20 via a plurality of second fuel inlets 130, so that the second fuel component 60 can also be injected through the injection wall 112 into the annular combustion chamber 34.

[0087] The conveying devices 68 and 114 are designed in the form of pumping devices, specifically in the form of turbopumps.

[0088] In one embodiment, the inner nozzle wall 24 and the outer nozzle wall 28 are made of a ceramic material.

[0089] In another embodiment of a thrust chamber device 10, the inner nozzle wall 24 and the outer nozzle wall 28 are made of a metallic material.

[0090] In another embodiment of the thrust chamber device 10, the inner nozzle wall 24 and the outer nozzle wall 28 are made of a ceramic material and a metallic material, respectively.

[0091] The inner nozzle wall 24 exhibits both in the Figure 2 the illustrated embodiment of the thrust chamber device 10 as well as in the Figures 3 to 5 The illustrated embodiments have the shape of a hyperboloid of revolution. This also applies accordingly to the inner shear space surface 26 and the outer shear space surface 36.

[0092] In all embodiments of the shear chamber device 10, the inner shear chamber surface 26 is convexly curved or substantially convexly curved, pointing towards the outer shear chamber surface 36.

[0093] Furthermore, in all described embodiments of shear chamber devices 10, the outer shear chamber surface 36 is convexly curved or substantially convexly curved in the direction of the inner shear chamber surface 26.

[0094] Furthermore, in all embodiments of shear chamber devices 10, the shear chambers 12 are rotationally symmetrical with respect to their longitudinal axis 30. This also applies accordingly to sections 14, 16 and 18.

[0095] Furthermore, in all embodiments of shear chamber devices 10, the outer shear chamber surface 36 and the inner shear chamber surface 26 are designed to be rotationally symmetrical with respect to the longitudinal axis 30.

[0096] The injection wall 112 delimits the thrust chamber 20 at the first end 42 of the first section 14. The injection wall 112 connects the inner nozzle wall 24 and the outer nozzle wall 28. As already explained, the majority of the first fuel inlets 110 and the majority of the second fuel inlets 130 are arranged in, formed, or penetrate the injection wall 112.

[0097] The ring-shaped injection wall 112 closes off the ring-shaped combustion chamber 34 at the first end 42.

[0098] The injection head 128 of the thrust chamber device 10 comprises the injection wall 112.

[0099] The majority of first fuel inlets 110 and the majority of second fuel inlets 130 are designed in the form of channels which have channel openings 132 pointing into the annular combustion chamber 34.

[0100] The majority of first fuel inlets 110 define first fuel inlet longitudinal axes 134, which point into the first section 14 in a direction parallel or substantially parallel to tangents to the inner thrust chamber surface 26 or the outer thrust chamber surface 36.

[0101] Furthermore, the thrust chamber device 10 comprises a first injection device 136 for injecting the first propellant component 56 into the thrust chamber 20 through the plurality of first propellant inlets 110. The first injection device 136 comprises the first conveying device 68.

[0102] The majority of second fuel inlets 130 define second fuel inlet longitudinal axes 138 which point into the first section 14 in a direction parallel or substantially parallel to tangents to the inner thrust chamber surface 26 or the outer thrust chamber surface 36.

[0103] The thrust chamber device 10 further comprises a second injection device 140 for injecting the second propellant component 60 into the thrust chamber 20 through the plurality of second propellant inlets 130. The second injection device 140 further comprises the second conveying device 114.

[0104] In Figure 3 A further embodiment of a thrust chamber device designated overall by reference numeral 10 is shown schematically.

[0105] In this embodiment, parallel regenerative cooling of the nozzle walls 24 and 28 with their regenerative cooling device 50 is possible.

[0106] As in the exemplary embodiment of the Figure 2The first conveying unit 68 is fluidly connected on the pressure side to the ring distributor 76 via the connecting line 74. The outer coolant channel inlets 80 of the outer coolant channels 78 are fluidly connected to the ring distributor 76 and extend through the outer nozzle wall 28 to the first end 42.

[0107] The pressure side 72 is fluidly connected to the collector 92 via a connecting line 142. Inner coolant channel inlets 86 of the inner coolant channels 88 are fluidly connected to the collector 92. The inner coolant channels 88 also extend to the first end 42 in the inner nozzle wall 24. Outer coolant channel outlets 82 are fluidly connected to inner coolant channel inlets 94 via the connecting line 84. The inner coolant channel outlets 94 are fluidly connected to the fluid inlet 98 of the drive unit 100 via connecting lines 144. Furthermore, the connecting line 106 fluidly connects the fluid outlet 102 to the distribution chamber 108.

[0108] The in Figure 3The regenerative cooling device 50, schematically depicted and described above, enables the coolant 52 to be conveyed by the first conveying device 68 in parallel, i.e., partially through the outer coolant channels 78 and simultaneously partially through the inner coolant channels 88. Such a parallel cooling mode for operating the cooling device 50 is particularly advantageous in a steady-state or load mode of the thrust chamber device.

[0109] In the load mode of the thrust chamber device 10, the nozzle walls 24 and 28 are sufficiently heated so that the coolant 52 is sufficiently heated when flowing through only the inner nozzle wall 24 or only the outer nozzle wall 28 in order to drive the drive device 100 in the desired manner.

[0110] Another embodiment of a thrust chamber device 10 is shown schematically in Figure 4It is shown in its operating principle of the regenerative cooling device 50 of the in Figure 3 similar to the illustrated embodiment.

[0111] At the in Figure 4 In the illustrated embodiment, outer coolant channel inlets 80 are fluidly connected to the pressure side 72 of the first conveying device 68 and direct the coolant 52 from the first end 42 to the ring distributor 76, which is fluidly connected to outer coolant channel outlets 82. The ring distributor 76, which in this embodiment functions as a collector, is fluidly connected to the fluid inlet 98 of the drive unit 100 via a connecting line 146. Furthermore, inner coolant channel inlets 86 are fluidly connected to the pressure side 72 via a connecting line 148.

[0112] The inner coolant channel outlets 94 are fluidly connected to the collector 92.

[0113] The collector 92 is in turn fluidly connected to the fluid inlet 98 either directly via the connecting line 96 or flows into the connecting line 146 in front of it.

[0114] The fluid outlet 102 of the drive unit 100 is connected to the distribution chamber 108 via connecting lines 106.

[0115] In the exemplary embodiment of the Figure 4 Parallel cooling of the nozzle walls 24 and 28 is also possible. The coolant 52 is distributed by the pumping device 68 to the inner and outer coolant channels 88 and 78 respectively, and thus flows either through one or the other coolant channel 78 or 88. In contrast to the embodiment of the Figure 3 In this embodiment, the cooling process takes place... Figure 4 starting from the first end 42 in a distal direction towards the third section 18. In the embodiment of the Figure 3The coolant 52 is conveyed towards the first end 42 through the nozzle walls 24 and 28 for cooling.

[0116] In Figure 5 A further embodiment of a thrust chamber device 10 is shown schematically. Its function is similar to the embodiment of the Figure 2 similar and enables a serial cooling mode of the regenerative cooling unit 50.

[0117] In this embodiment, the pressure side 72 is fluidly connected to the collector 92 in the region of the distal end 90 of the insert body 22 via a connecting line 150. Internal coolant channel inlets 86 are directly fluidly connected to the collector 92 and extend through the inner nozzle wall 24 towards the first end 42.

[0118] Connecting lines 84 in the region of the first end 42 connect inner coolant channel outlets 94 fluidly with outer coolant channel inlets 80 in the region of the first end 42. The outer coolant channels 78 extend through the outer nozzle wall 28 to the ring distributor 76, which in this embodiment again assumes the function of a collector.

[0119] The ring distributor 76 is fluidly connected to the fluid inlet 98 of the drive unit 100 via a connecting line 146. The fluid outlet 102 of the drive unit 100 is in turn fluidly connected to the distributor chamber 108 via connecting lines 106.

[0120] With the exemplary embodiment of Figure 5 Serial cooling of the nozzle walls 24 and 28 can be implemented. Unlike the embodiment of the Figure 2, where the coolant 52 first flows through the outer nozzle wall 28 and then through the inner nozzle wall 24, cooling takes place in the exemplary embodiment of the Figure 5 Quite the opposite. The coolant 52 is first passed through the inner nozzle wall 24 and then through the outer nozzle wall 28.

[0121] In Figure 7 A coolant flow switching device 152 is shown schematically. The embodiments of the thrust chamber devices 10 described above can optionally include such a coolant flow switching device 152. If this is provided, an inlet 154 of the device is fluidly connected to the pressure side 72 of the first conveying device 68 via a connecting line 156.

[0122] The coolant flow switching device 152 comprises, in the Figure 7In the illustrated embodiment, three outputs 158, 160, and 162 are connected to the inner and outer coolant channels 88 and 78, respectively, such that the coolant flow switching device 152 can be used, for example, to switch between a parallel cooling mode and a series cooling mode. This makes it possible, in particular, by appropriately connecting the outputs 158, 160, and 162 to the coolant channels 78 and 88, to utilize both the functions described above in connection with the embodiments of the Figures 2 and 3 described cooling modes, i.e., on the one hand serial as in Figure 2 or parallel as in Figure 3 to realize this. Alternatively, the implementation examples of the Figures 4 and 5with the coolant flow switching device 152 accordingly realized on a single thrust chamber device 10, so that, for example, when starting the thrust chamber device 10, the regenerative cooling device 50 is operated with a serial cooling mode, and during load operation or steady-state operation with a parallel cooling mode.

[0123] As schematically in Figure 8 As shown, an engine 166, for example for a missile 168 or an aircraft, is equipped with a thrust chamber device 10. The thrust chamber device 10 is designed in the form of a thrust chamber device 10 according to one of the embodiments described above.

[0124] Figure 9 Figure 1 schematically shows an embodiment of a missile 168 or a flying device comprising an engine 166 with a thrust chamber device 10. In one embodiment, the missile 168 is designed in the form of a rocket.

[0125] The described embodiments of thrust chamber devices 10 enable efficient cooling of the nozzle walls 24 and 28. Furthermore, the special shape of the inner and outer thrust chamber surfaces 26 and 36 significantly reduces the overall length of each thrust chamber 12 compared to conventional thrust chamber devices with a cylindrical thrust chamber or a cylindrical combustion chamber to which a Laval nozzle is attached to form the narrowest point. This makes it possible to design thrust chamber devices 10 with a significantly reduced weight compared to known thrust chamber devices. Thus, each of the thrust chamber devices 10 described above can deliver a significantly higher payload, for example, into near-Earth space, while providing the same performance as a conventional thrust chamber device. Overall, a thrust chamber device 10 can be operated very efficiently in this way. Reference symbol list

[0126] 10 Thrust chamber device 12 Thrust chamber 14 First section 16 Second section 18 Third section 20 Thrust chamber 22 Insert body 24 Inner nozzle wall 26 Inner thrust chamber surface 28 Outer nozzle wall 30 Longitudinal axis 32 Tangent 34 Annular combustion chamber 36 Outer thrust chamber surface 38 Cross-sectional area 40 Spacing 42 First end 44 Transition 46 Narrowest point 48 Nozzle outlet 50 Cooling device 52 Coolant 54 First fuel reservoir 56 First fuel component 58 Second fuel reservoir 60 Second fuel component 62 Fuel 64 Oxidizer 66 Connecting line 68 First conveying device 70 Suction side 72 Pressure side 74 Connecting line 76 Ring distributor 78 Outer coolant channel 80 Outer coolant channel inlet 82 Outer coolant channel outlet 84 Connecting line 86 Inner coolant channel inlet 88 Inner coolant channel 90 Distal end 92 Collector 94 Inner coolant channel outlet 96 Connecting line 98 Fluid inlet 100 Drive unit 102 Fluid outlet 104 Turbine 106 Connecting line 108 Distribution chamber 110 FirstFuel inlet 112 Injection wall 114 Secondary delivery device 116 Drive shaft 118 Connecting line 120 Suction side 122 Connecting line 124 Pressure side 126 Distribution chamber 128 Injection head 130 Secondary fuel inlets 132 Channel opening 134 First fuel inlet longitudinal axis 136 First injection device 138 Second fuel inlet longitudinal axis 140 Second injection device 142 Connecting line 144 Connecting line 146 Connecting line 148 Connecting line 150 Connecting line 152 Coolant flow switching device 154 Inlet 156 Connecting line 158 Outlet 160 Outlet 162 Outlet 164 Valve assembly 166 Engine 168 Missile

Claims

1. Thrust chamber device (10), comprising a thrust chamber (12) with a thrust space (20) that has a first portion (14), a second portion (16) adjoining the first portion (14), and a third portion (18) adjoining the second portion (16), wherein the thrust space (20) is delimited in all three portions (14, 16, 18) by an outer nozzle wall (28) with an outer thrust space surface (36), which outer thrust space surface (36) tapers in the first and second portion (14, 16) toward the third portion (18) and in the third portion (18) expands away from the second portion (16), wherein a narrowest point (46) is formed at the transition from the second portion (16) to the third portion (18), wherein the first portion (14) is delimited by an inner nozzle wall (24) with an inner thrust space surface (26), which tapers toward the second portion (16), and wherein formed between the inner thrust space surface (26) and the outer thrust space surface (36) is an annular combustion space (34), which extends over the first portion (14), characterized in that the thrust chamber device (10) comprises a regenerative cooling unit (50) for cooling the inner nozzle wall (24) and the outer nozzle wall (28) with a coolant (52), in order to fully use the heat or enthalpy absorbed when cooling the two nozzle walls (24, 28) for driving a drive unit (100), with which one or more conveying units (68, 114) for propellant components (56, 58) to be combusted in the thrust space (20) can be driven.

2. Thrust chamber device in accordance with Claim 1, characterized in that a) the regenerative cooling unit (50) comprises a plurality of inner coolant channels (88) in the inner nozzle wall (24) and a plurality of outer coolant channels (78) in the outer nozzle wall (28) and in that the plurality of inner coolant channels (88) and the plurality of outer coolant channels (78) are configured to be flowed through by the coolant (52) and / or b) the inner nozzle wall (24) and / or the outer nozzle wall (28) are made of a ceramic and / or metallic material and / or c) the inner thrust space surface (26) and / or the outer thrust space surface (36) at least partially have the form of a hyperboloid of revolution or have a continuously concave longitudinal sectional line and / or d) the inner thrust space surface (26) is of convexly curved or substantially convexly curved configuration pointing in the direction toward the outer thrust space surface (36) and / or e) the outer thrust space surface (36) is of convexly curved or substantially convexly curved configuration pointing in the direction toward the inner thrust space surface (26).

3. Thrust chamber device in accordance with any one of the preceding Claims, characterized in that a) the thrust chamber (12) defines a longitudinal axis (30) and in that the thrust chamber (12), in particular the first portion (14) and / or the second portion (16) and / or the third portion (18), are of rotationally symmetrical configuration relative to the longitudinal axis (30), wherein, in particular, the outer thrust space surface (36) and / or the inner thrust space surface (26) are of rotationally symmetrical configuration relative to the longitudinal axis (30), and / or b) the annular combustion space (34) has a constant or substantially constant cross sectional area.

4. Thrust chamber device in accordance with any one of the preceding Claims, characterized in that the thrust chamber device (10) comprises a plurality of first propellant inlets (110) for a first propellant component (56) and a plurality of second propellant inlets (130) for a second propellant component (60), wherein, in particular, the first portion (14) of the thrust space (20) is delimited on an end (42) pointing away from the second portion (16) by an injection wall (112), which connects the inner nozzle wall (24) and the outer nozzle wall (28) to one another, and wherein the plurality of first propellant inlets (110) and the plurality of second propellant inlets (130) are arranged or formed in the injection wall (112), wherein, further in particular, a) the injection wall (112) is of annular or rotationally symmetrical or hyperboloid-like configuration for closing the ring-shaped annular combustion space (34) and / or b) the thrust chamber device (10) comprises an injection head (128) and wherein the injection head (128) comprises the injection wall (112).

5. Thrust chamber device in accordance with Claim 4, characterized in that a) the plurality of first propellant inlets (110) and the plurality of second propellant inlets (130) are configured in the form of channels, which have channel openings (132) pointing into the annular combustion space (34), and / or b) the plurality of first propellant inlets (110) define first propellant inlet longitudinal axes (134) and in that the first propellant inlet longitudinal axes (134) point into the first portion (14) in a direction parallel or substantially parallel to tangents to the inner thrust space surface (26) and / or the outer thrust space surface (36) and / or c) the thrust chamber device comprises a first injection unit (136) for injecting the at least one first propellant component (56) into the thrust space (20) through the plurality of first propellant inlets (110), wherein, in particular, the thrust chamber device (10) comprises a first propellant store (54) for a first propellant component (56) and wherein the first injection unit (136) comprises a first conveying unit (68) for conveying the at least one first propellant component (56) from the first propellant store (54) through the plurality of first propellant inlets (110) into the thrust space (20), wherein, further in particular, the first conveying unit (68) has a suction side (70) and a pressure side (72), wherein the suction side (70) is fluidically connected to the first propellant store (54), and wherein the pressure side (72) is fluidically connected to the outer coolant channels (78) and / or the inner coolant channels (88), and / or d) the plurality of second propellant inlets (130) define second propellant inlet longitudinal axes (138) and in that the second propellant inlet longitudinal axes (138) point into the first portion (14) in a direction parallel or substantially parallel to tangents to the inner thrust space surface (26) and / or the outer thrust space surface (36) and / or e) the thrust chamber device comprises a second injection unit (140) for injecting the at least one second propellant component (60) into the thrust space (20) through the plurality of second propellant inlets (130), wherein, in particular, the thrust chamber device (10) comprises at least one second propellant store (58) for at least one second propellant component (60) and wherein the second injection unit (140) comprises a second conveying unit (114) for conveying the at least one second propellant component (60) from the second propellant store (58) through the plurality of second propellant inlets (130) into the thrust space (20) and / or f) the first propellant component (56) forms the coolant (52) and / or g) the first propellant component (56) is a liquid fuel (62), in particular liquid hydrogen or liquid methane or liquefied natural gas, and / or h) the second propellant component (60) is a liquid oxidizer (64), in particular liquid oxygen.

6. Thrust chamber device in accordance with Claim 4 or 5, characterized in that the thrust chamber device comprises a first injection unit (136) for injecting the at least one first propellant component (56) into the thrust space (20) through the plurality of first propellant inlets (110), in that the thrust chamber device (10) comprises a first propellant store (54) for a first propellant component (56), in that the first injection unit (136) comprises a first conveying unit (68) for conveying the at least one first propellant component (56) from the first propellant store (54) through the plurality of first propellant inlets (110) into the thrust space (20), and in that the thrust chamber device (10) comprises a drive unit (100) for driving the first conveying unit (68) and / or the second conveying unit (114), wherein, in particular, the drive unit (100) a) is configured in the form of a turbine (104), in particular a gas turbine, and / or b) comprises a fluid inlet (98) and a fluid outlet (102) and wherein the fluid outlet (102) is fluidically connected to the first propellant inlets (110).

7. Thrust chamber device in accordance with any one of Claims 2 to 6, characterized in that the outer coolant channels (78) comprise outer coolant channel inlets (80) and outer coolant channel outlets (82) and in that the inner coolant channels (88) comprise inner coolant channel inlets (86) and inner coolant channel outlets (94), wherein, in particular, the pressure side (72) of the first conveying unit (68) a) is fluidically connected to the outer coolant channel inlets (80), wherein the outer coolant channel outlets (82) are fluidically connected to the inner coolant channel inlets (86), and wherein the inner coolant channel outlets (94) are fluidically connected to the fluid inlet (98) of the drive unit (100) or b) is fluidically connected to the outer coolant channel inlets (80) and the inner coolant channel inlets (86), and wherein the outer coolant channel outlets (82) and the inner coolant channel outlets (94) are fluidically connected to the fluid inlet (98) of the drive unit (100) or c) is fluidically connected to the inner coolant channel inlets (86), wherein the inner coolant channel outlets (82) are fluidically connected to the outer coolant channel inlets (80), and wherein the outer coolant channel outlets (82) are fluidically connected to the fluid inlet (98) of the drive unit (100).

8. Thrust chamber device in accordance with Claim 7, characterized in that the thrust chamber device (10) comprises a coolant flow switching unit (152) for selectively switching a coolant flow in parallel or serially through the inner coolant channels (88) and the outer coolant channels (78), wherein, in particular, the coolant flow switching unit (152) comprises a valve unit (164).

9. Engine (166), in particular for a flying object (168) or an aircraft, comprising a thrust chamber device (10) in accordance with any one of the preceding Claims.

10. Flying object (168) or aircraft, comprising a first propellant store (54) for at least one first propellant component (56), a second propellant store (58) for at least one second propellant component (60), and an engine (166) in accordance with Claim 9.

11. Method for operating a thrust chamber device (10) in accordance with any one of Claims 1 to 8, characterized in that the inner nozzle wall (24) and the outer nozzle wall (28) are regeneratively cooled with a coolant (52), in order to fully use the heat or enthalpy absorbed by the coolant (52) when cooling the two nozzle walls (24, 28) for driving a drive unit (100), with which one or more conveying units (68, 114) for propellant components (56, 58) to be combusted in the thrust space (20) can be driven.

12. Method in accordance with Claim 11, characterized in that the first propellant component (56) is used as coolant (52).

13. Method in accordance with Claim 11 or 12, characterized in that the coolant (52) a) is conducted in parallel through the outer nozzle wall (28) and the inner nozzle wall (24), in particular when the thrust chamber device (10) is in a stable operating mode or in a load mode, or b) is conducted first through the inner nozzle wall (24) and then through the outer nozzle wall (28) or first through the outer nozzle wall (28) and then through the inner nozzle wall (24).

14. Method in accordance with any one of Claims 11 to 13, characterized in that a cooling mode is changed when changing an operating mode of the thrust chamber device (10), wherein, in particular, in a start mode of the thrust chamber device (10) a serial cooling mode is performed in accordance with Claim 13 and wherein in a load mode of the thrust chamber device (10) a parallel cooling mode is performed in accordance with Claim 13.

15. Method in accordance with any one of Claims 11 to 14, characterized in that the coolant (52) a) flows through the drive unit (100) before or after flowing through the inner nozzle wall (24) and the outer nozzle wall (28) and / or b) is injected through the first propellant inlets (110) into the thrust chamber (12) after flowing through the inner nozzle wall (24) and / or the outer nozzle wall (28).