Modular engine for missiles, procedures and use of a modular engine

Abstract

Modular engine (1) for missiles, comprising: an engine inlet (10); a gas turbine (12); a payload turbine (14); at least one propulsor (20) and a nozzle assembly (16), wherein an outlet of the engine inlet (10) is connected to an inlet of the gas turbine (12), and wherein an outlet of the gas turbine (12) is selectively connectable to an inlet of the nozzle assembly (16) and / or to an inlet of the payload turbine (14) for activating and deactivating the payload turbine (14) for driving the at least one propulsor (20), characterized in that the engine (1) is designed to deactivate the payload turbine by jettisoning the payload turbine (14) and / or the at least one propulsor (20).

Description

The present invention relates to a modular engine for missiles, a method for operating a modular engine and a use of the modular engine. Conventional engines employ a specific propulsion system depending on the aircraft being powered. Slow-moving aircraft, in particular, often utilize propeller propulsion, as this allows for high thrust efficiency and thus a long flight range, although the achievable airspeed is limited. A propeller engine can, for example, be designed as a turboprop engine and includes a propeller, an engine intake, a gas turbine with a payload turbine to drive the propeller, and an exhaust nozzle. To achieve higher flight speeds, turbojet engines are typically used as a second propulsion method, specifically turbojet engines, which comprise an engine intake, a gas turbine, and a thrust nozzle. However, these generally exhibit lower thrust efficiency compared to propeller-driven engines, particularly during low-speed flight phases. High-speed applications are especially relevant for military purposes, as they reduce reaction time during flight maneuvers and increase survivability. For a propulsion system based on the prior art, the type of propulsion is predetermined and cannot be changed later, for example, during flight. Consequently, when designing the propulsion system for a missile, a choice must be made between achieving a high flight speed or achieving high propulsion efficiency and thus a long flight range. In this context, the differences between a propeller drive and a turbojet drive are particularly significant for small and / or unmanned aircraft, such as cruise missiles. This is a major disadvantage, especially for missions with varying requirements regarding flight range and flight speed in individual mission segments, such as complex military missions. From DE 29 71 303 A1, a propulsion system for aircraft has become known, comprising two-circuit turbojet engines which have a hot primary circuit and a cold secondary circuit for generating forward thrust, wherein the secondary circuit can be driven by the primary circuit in the manner of a blower or propeller, wherein a bypass with adjustable reversing flaps is provided for the gas jet of the primary circuit to selectively bypass the drive of the blower, and the blower blades are arranged in conjunction with an adjustment device to be selectively adjustable into a low-drag position. DE 10 2013 209 538 A1 relates to a hybrid drive for a power-driven aircraft, comprising a primary shaft non-rotatably coupled or connectable to a gas turbine and a secondary shaft non-rotatably coupled or connectable to an electric machine, wherein a turbine wheel is also non-rotatably coupled or connectable to the secondary shaft, by means of which an exhaust gas flow generated during operation of the gas turbine can be used to drive the secondary shaft. The hybrid drive includes a propeller to which a torque from the primary shaft and / or a torque from the secondary shaft can be transmitted. Based on the aforementioned disadvantages of the prior art, the present invention aims to provide a modular propulsion device for missiles, a method for operating a modular engine, and a use of the modular engine, so that both energy-efficient propulsion generation for a long flight range at low flight speeds and high flight speeds can be achieved temporarily during flight segments. According to a first aspect of the present invention, a modular missile engine is provided, comprising an engine inlet, a gas turbine, a payload turbine, at least one propulsor, and a nozzle assembly, wherein an outlet of the engine inlet is connected to an inlet of the gas turbine, and wherein an outlet of the gas turbine is selectively connectable to an inlet of the nozzle assembly and / or to an inlet of the payload turbine for activating and deactivating the payload turbine to drive the at least one propulsor. An outlet of the payload turbine can be connected to the inlet of the nozzle assembly, particularly when the outlet of the gas turbine is connected to the inlet of the payload turbine to drive the at least one propulsor. Alternatively, the payload turbine can have an exhaust nozzle in addition to the nozzle assembly of the modular engine.The connections between the individual components are designed in such a way that, in particular, further components may be present between the outlet of the gas turbine and the inlet of the nozzle assembly and / or the outlet of the gas turbine and the inlet of the payload turbine. An advantage of the invention is that the propulsion device exhibits modularity with regard to its drive type and is designed for propulsion with at least one propulsor as well as with a nozzle assembly. The payload turbine can be activated and deactivated to drive the at least one propulsor, so that thrust generation is energy-efficiently achieved by the at least one propulsor or, for high flight speeds, by the fluid jet through the nozzle assembly. Furthermore, the payload turbine for driving the at least one propulsor can be retrofitted to an existing turbojet engine, since the structural and design modifications are essentially located downstream of the gas turbine outlet.Thus, a large part of the turbojet drive, in particular the gas turbine, remains identical in design, which leads to significant cost reductions and reduced development effort when retrofitting the solution according to the invention to existing turbojet engines. In a preferred embodiment, the modular engine additionally comprises a drive train, wherein the drive train is configured for power transmission from the payload turbine to the at least one propulsor, and wherein the power transmission is mechanical and / or electrical. The drive train can be configured as a direct connection directly to the payload turbine or, for example, extend between the payload turbine and the at least one propulsor in the form of a shaft and / or electrical lines, and may include further components for power transmission. Thus, the integration of the modular engine on the missile can be flexibly adapted depending on the application. Preferably, power transmission via the drivetrain is mechanical, with the load transmission from the payload turbine to the at least one propulsor being configured as a direct drive, a drive via a shaft, and / or a drive via a gearbox. Mechanical power transmission is particularly advantageous for high power applications due to its low transmission losses. Furthermore, mechanical power transmission is especially suitable for unbranched or minimally branched drivetrains. In both direct and shaft drives, the at least one propulsor and the payload turbine rotate at the same speed. To adjust the rotational speeds between the payload turbine and the propeller, a gearbox can be used between the two components. In a preferred embodiment, power transmission via the drive train is electrical, wherein the drive train comprises a generator mechanically connected to the payload turbine and an electric drive, such as an electric motor, mechanically connected to the at least one propulsor. Particularly preferably, an electrical energy storage device is arranged between the generator and the electric motor, which is designed for intermediate storage of the electrical energy generated by the generator. Electrical power transmission is particularly suitable for small installation spaces and branched drive trains. Furthermore, an electric drive train facilitates the distribution of power to multiple propulsors. The electrical storage device is, in particular, designed to be switchable on and off.Thus, the intermediate storage unit can be used for energy absorption or release, and can temporarily store excess energy during different drive phases, or be used to drive at least one propulsor, or to provide additional support for the drive of at least one propulsor. The provision of the electrical energy storage unit allows for the gas turbine to be temporarily switched off and for purely electric propulsion using the stored energy of the intermediate storage unit. The at least one propulsor has, in particular, a propeller. Furthermore, the at least one propulsor for propelling the missile can be configured differently. In another preferred embodiment, the at least one propulsor can have a ducted fan, fan, open rotor, rotor, or the like. The present invention is not limited to one embodiment of the propulsor, so that the modular propulsion device can also have different propulsors simultaneously. The propeller can have folding rotor blades. The adjustment of the rotor blades is achieved primarily by a rotor adjustment mechanism, which is designed such that the rotor blades can be folded in and out radially or, alternatively, axially. In the extended state, the propeller performs work on the surrounding fluid to generate thrust, while in the folded state, the rotor blades experience low drag and are unloaded. Furthermore, the propeller can also be designed as a centrifugal folding propeller that folds into the fuselage of the missile. The rotor blades are extended when the payload turbine is activated and folded when the payload turbine is deactivated. In a preferred embodiment, the modular engine additionally features an adjustment mechanism, the adjustment mechanism being configured to selectively direct the fluid flow from the gas turbine outlet to the payload turbine inlet or to the nozzle assembly inlet for activating or deactivating the payload turbine. The adjustment mechanism is, in particular, arranged between the gas turbine outlet and the nozzle assembly inlet and / or between the gas turbine outlet and the payload turbine inlet. In the case where the fluid flow is directed from the gas turbine outlet to the payload turbine inlet for flow through the payload turbine, the payload turbine outlet can be connected to the nozzle assembly inlet.Thus, the fluid flow can first pass through the payload turbine and subsequently the nozzle assembly in series, whereby in the case of an activated payload turbine, the fluid flow can be expanded through the nozzle assembly to ambient conditions. Preferably, the adjustment mechanism includes a valve or flap for redirecting the fluid flow exiting the gas turbine outlet. Thus, the redirection of the fluid flow can be set by specifying the valve or flap position, and multiple switching between positions is possible. In a preferred embodiment, the nozzle assembly has a variable surface geometry. The surface geometry can be adjusted, in particular, depending on the different fluid states at the inlet of the nozzle assembly and the ambient conditions at the outlet of the nozzle assembly, so that the conversion of fluid energy into a high jet velocity is as complete as possible, or the fluid can expand to ambient conditions. This is particularly advantageous when the nozzle assembly is used for thrust generation in one drive phase and for expansion of the fluid flow after it has passed through the payload turbine in a second drive phase. The engine can be configured to activate the payload turbine by redirecting the fluid flow and allowing the fluid flow from the gas turbine outlet to pass through the payload turbine. When the fluid flows through the payload turbine, the fluid performs work on the turbine blades, and power can be drawn through the payload turbine to drive at least one propulsor, thus activating the payload turbine. The fluid flow is then expanded through the nozzle assembly or an exhaust nozzle. In a preferred embodiment, the engine is configured to deactivate the payload turbine by diverting the fluid flow from the payload turbine to the nozzle assembly, where the fluid flow from the gas turbine outlet passes through the nozzle assembly, thus preventing the fluid flow from reaching the payload turbine. If the fluid flow does not pass through the payload turbine, the fluid performs no work on the payload turbine and no power can be absorbed by the payload turbine. Consequently, the drive of the at least one propulsor remains inactive, and the payload turbine is deactivated. In this case, the fluid state at the nozzle assembly inlet corresponds to the state at the gas turbine outlet and is converted into jet velocity for thrust generation as it flows through the nozzle assembly. According to the invention, the engine is designed to deactivate the payload turbine by jettisoning the payload turbine and / or the at least one propulsor. If propulsion via the payload turbine and the at least one propulsor is no longer required on a flight mission, jettisoning the payload turbine, preferably together with the propeller and especially preferably with the drive train, reduces the weight of the propulsion system and thus increases the flight range. Switching between the propulsion modes is then no longer possible. In a preferred embodiment, the engine additionally comprises a control unit configured to control the activation and deactivation of the payload turbine. This can be achieved, in particular, by controlling the adjustment mechanism. Thus, switching between different propulsion modes is possible during operation. The control unit is designed such that the activation and deactivation of the payload turbine can be controlled by direct input from the pilot in the missile, remotely via a control command, or by a pre-programmed sequence. Furthermore, the control unit can, in particular, adapt the variable surface geometry of the respective propulsion phase. Preferably, the rotor pitch is also controlled by the control unit. A second aspect of the present invention relates to a method for operating a modular engine with a first and at least one second drive phase of the engine, wherein in the first drive phase the payload turbine is activated to drive the at least one propulsor by passing the fluid flow from the gas turbine outlet through the payload turbine, and in the at least second drive phase the payload turbine is deactivated and the fluid flow from the gas turbine outlet is expelled through the nozzle assembly. The different drive phases can be used according to the respective requirements in the individual segments of a flight mission. The designation of the first drive phase and the at least second drive phase does not restrict the order of their use. In a preferred embodiment of the method, in the first drive phase, the thrust is generated primarily by the at least one propulsor, wherein the at least one propulsor is driven by the payload turbine, and in at least a second drive phase, the thrust is generated primarily by the expulsion of a fluid jet from the nozzle assembly. Thus, the first drive phase and the at least second drive phase have different drive types. Optionally, the engine switches between the first and second propulsion phases at least once during operation. This allows the propulsion system to be adapted to the specific requirements of each segment of a flight mission. The sequence of the first and second propulsion phases can be varied as desired. The payload turbine and / or the at least one propulsor can be jettisoned, preferably jettisoned, from the modular engine during the transition from the first to the second propulsion phase. This is particularly advantageous when thrust generation by the propeller is no longer required for the remainder of the flight mission. This makes the propulsion system lighter and allows for greater flight ranges. A third aspect of the present invention relates to the use of a modular engine according to the first aspect of the invention for carrying out the method according to the second aspect of the invention. In the following, exemplary embodiments of the modular drive system according to the invention are schematically explained with reference to the attached figures. Figure 1A shows a schematic sectional view in side view of a modular engine according to a first exemplary embodiment of the present invention; Figure 1B shows a schematic sectional view in side view of a modular engine according to a second exemplary embodiment of the present invention; Figure 2 shows a schematic sectional view in side view of a modular engine according to a third exemplary embodiment of the present invention; Figure 3 shows a schematic sectional view in side view of a modular engine according to a fourth exemplary embodiment of the present invention; and Figure 4 shows a schematic representation of a method according to a preferred embodiment of the present invention. As shown in Fig. 1, a modular engine 1 for a missile according to the present invention comprises an engine inlet 10, a gas turbine 12, a payload turbine 14, at least one propulsor 20 and a nozzle assembly 16, wherein an outlet of the engine inlet 10 is connected to the inlet of the gas turbine 12 and wherein an outlet of the gas turbine 12 can be selectively connected to an inlet of the nozzle assembly 16 and / or to an inlet of the payload turbine 14 for activating and deactivating the payload turbine 14 for driving the at least one propulsor 20. Fig. 1A shows a first preferred embodiment of the present invention, wherein the outlet of the gas turbine 12 is connected to the inlet of the payload turbine 14, and wherein the outlet of the payload turbine 14 is connected to the inlet of the nozzle assembly 16. The fluid flow at the outlet of the gas turbine 12 thus first flows through the payload turbine 14 and then through the nozzle assembly 16, where the fluid flow expands to ambient conditions. Alternatively, the outlet of the gas turbine 12 is connected to the inlet of the nozzle assembly 16, and the fluid flow at the outlet of the gas turbine 12 flows through the nozzle assembly 16 without first flowing through the payload turbine 14, whereby the fluid flow is converted into a fluid jet for propulsion generation. Fig. 1B shows a second preferred embodiment of the present invention, wherein the outlet of the gas turbine 12 is connected to the inlet of the payload turbine 14 and the inlet of the nozzle assembly 16, while the outlet of the payload turbine 14 is not connected to the inlet of the nozzle assembly 16. Thus, the fluid flow at the outlet of the gas turbine 12 flows either through the payload turbine 14 to drive the at least one propulsor 20 or through the nozzle assembly to generate a fluid jet. In this case, the payload turbine has its own exhaust nozzle 145, through which the fluid flow is expanded to ambient conditions. In the embodiments shown in Fig. 1, the at least one propulsor 20 is directly connected to the payload turbine 14, so that the at least one propulsor 20 and the payload turbine 14 have a common rotational speed. Fig. 2 shows a third preferred embodiment of the present invention, wherein the at least one propulsor 20 is connected to the payload turbine 14 via a drive train 22. Power transmission from the payload turbine 14 to the at least one propulsor 20 can be mechanical and / or electrical. In an embodiment where power transmission via the drive train 22 is mechanical, the drive train 22 can, for example, be designed as a shaft. Furthermore, in a preferred embodiment, the drive train 22 can include a gearbox 24, so that the rotational speed of the payload turbine 14 is reduced relative to that of the at least one propulsor 20. Thus, the payload turbine 14 can operate at higher speeds than the propulsor 20. In a preferred embodiment, power transmission via the drive train 22 is electrical, wherein the drive train 22 comprises a generator mechanically connected to the payload turbine 14 and an electric motor mechanically connected to the at least one propulsor 20. Preferably, an electrical energy storage device 26, designed for intermediate storage of the electrical energy generated by the generator, is additionally arranged between the generator and the electric motor. The drive train 22 includes, for example, electrical conductors. In a preferred embodiment, a gearbox can be arranged between the payload turbine 14 and the mechanically connected generator in the case of electrical power transmission via the drive train 22, and / or a gearbox can be arranged between the at least one propulsor 20 and the mechanically connected generator.Thus, the generator can be operated at a different speed, in particular a higher speed, than the payload turbine 14, and can therefore be designed more compactly. Similarly, the electric motor can be operated at a different, in particular a higher speed, than the at least one propulsor 20, and can therefore be designed more compactly. According to the invention, the power transmission from the payload turbine 14 to the at least one propulsor 20, as described above, can also be provided in conjunction with an arrangement of inlet 10, gas turbine 12, payload turbine 14 and nozzle assembly 16 according to the embodiment shown in Fig. 1B. The propulsor 20 shown in Figures 1-4, in the form of a propeller, represents a preferred embodiment of the propulsor 20. In a preferred embodiment, the propeller has foldable and extendable rotor blades, so that in the extended state the propeller performs work on the surrounding fluid to generate thrust, and in the folded state the rotor blades are free from resistance and load. Furthermore, in another preferred embodiment, the at least one propulsor 20 can have a ducted fan, open rotor, rotor, or the like. According to a preferred embodiment of the present invention, Fig. 3 additionally features an adjustment mechanism 18 between the outlet of the gas turbine 12 and the inlet of the payload turbine 14, as well as the inlet of the nozzle assembly 16. The adjustment mechanism 18 is configured to selectively direct the fluid flow from the outlet of the gas turbine 12 to the inlet of the payload turbine 14 or to the inlet of the nozzle assembly 16 for activating or deactivating the payload turbine 14. The present invention is not limited to the embodiment shown, so that the fluid flow can also be expanded to ambient conditions after passing through the payload turbine 14 via the nozzle assembly 16. In this case, the exhaust nozzles 145 are omitted. In a preferred embodiment, the adjustment mechanism 18 includes, in particular, a valve or flap for redirecting the fluid flow exiting the gas turbine 12 outlet. Consequently, the payload turbine 14 is activated when the adjustment mechanism 18 directs the fluid flow at the gas turbine 12 outlet into the payload turbine 14 inlet, thus causing the fluid to perform work at the payload turbine 14 and driving the at least one propulsor 20. Conversely, the payload turbine 14 is deactivated when the adjustment mechanism 18 directs the fluid flow at the gas turbine 14 outlet into the nozzle assembly 16 inlet, where it can be converted into jet velocity for propulsion generation. In this case, the payload turbine 14 is not subjected to fluid flow, and the at least one propulsor 20 is not driven. Fig. 4 shows an embodiment of the invention, wherein the modular engine 1 is configured to deactivate the payload turbine 14 by jettisoning the payload turbine 14 and / or the at least one propulsor 20 and / or the drive train. For example, the adjustment mechanism 18 can direct the fluid flow through the nozzle assembly 16 from the outlet of the gas turbine 12, so that the payload turbine 14 is no longer supplied with fluid. This reduces the weight of the propulsion device 1 and increases the flight range. Subsequently, propulsion is generated exclusively by the fluid jet from the nozzle assembly 16. Switching between the two propulsion modes is then no longer possible. The embodiments shown in Figures 1-4 can, in particular, have a variable surface geometry in the nozzle assembly 16. This is especially relevant when the nozzle assembly 16 is used both for generating thrust by the fluid jet and for expanding the fluid flow after passing through the payload turbine 14. In a preferred embodiment, the embodiments shown in Figures 1-4 can include a control unit, wherein the control unit is configured to control the activation and deactivation of the payload turbine 14. The control unit can, in particular, control the adjustment mechanism 18 and / or the folding and unfolding rotor blades and / or the jettisoning of the payload turbine 14, preferably with the at least one propulsor 20 and especially preferably together with the drive train 22. The control unit can be pre-programmed for pilot input or operated remotely. In a further aspect of the present invention, the modular engine 1 can, for example, be operated according to a method with a first and at least one second drive phase of the engine, wherein in the first drive phase the payload turbine 14 is activated to drive the at least one propulsor 20 by passing the fluid flow from the outlet of the gas turbine 12 through the payload turbine 14 and in the at least second drive phase the payload turbine 14 is deactivated and the fluid flow from the outlet of the gas turbine 12 is expelled through the nozzle assembly 16. For example, in the first propulsion phase, the thrust generation can be primarily carried out by the at least one propulsor 20, wherein the at least one propulsor 20 is driven by the payload turbine 14, and in at least a second propulsion phase, the thrust generation can be primarily carried out by the emission of a fluid jet from the nozzle device 16. Furthermore, the engine can switch between the first and second propulsion phases at least once during operation. For example, the missile can initially be propelled by the propulsor with the payload turbine activated to achieve the greatest possible flight range. Subsequently, the modular engine switches to the second propulsion phase, so that thrust is generated by the fluid jet from the nozzle assembly 16 and the payload turbine 14 is deactivated. This allows the missile to achieve higher flight speeds and makes it more agile in this flight segment. If propulsion via the at least one propulsor 20 driven by the payload turbine 14 is no longer required for the remainder of the flight mission, the payload turbine 14 and / or the at least one propulsor 20 can be jettisoned, preferably jettisoned, from the modular engine 1 during the transition from the first to the second propulsion phase. This reduces the weight of the modular engine 1 and increases its power-to-weight ratio. Reference symbol list: 1 Modular engine 10 Engine inlet 12 Gas turbine 14 Payload turbine 16 Nozzle assembly 18 Adjustment mechanism 20 Propeller 22 Drivetrain 24 Gearbox 26 Electrical storage 145 Exhaust nozzle

Claims

Modular engine (1) for missiles, comprising: an engine inlet (10); a gas turbine (12); a payload turbine (14); at least one propulsor (20) and a nozzle assembly (16), wherein an outlet of the engine inlet (10) is connected to an inlet of the gas turbine (12), and wherein an outlet of the gas turbine (12) is selectively connectable to an inlet of the nozzle assembly (16) and / or to an inlet of the payload turbine (14) for activating and deactivating the payload turbine (14) for driving the at least one propulsor (20), characterized in that the engine (1) is designed to deactivate the payload turbine by jettisoning the payload turbine (14) and / or the at least one propulsor (20). Modular engine (1) according to claim 1, further comprising a drive train (22), wherein the drive train (22) is designed for power transmission from the payload turbine (14) to the at least one propulsor (20), and wherein the power transmission is carried out mechanically and / or electrically. Modular propulsion system (1) according to claim 2, wherein the power transmission via the drive train (22) is mechanical, and wherein the load transmission from the payload turbine (14) to the at least one propulsor (20) is designed as a direct drive or a drive via a shaft and / or a drive via a gearbox (24). Modular propulsion system (1) according to one of claims 2 or 3, wherein the power transmission via the drive train (22) is electrical, wherein the drive train has a generator mechanically connected to the payload turbine (14) and an electric motor mechanically connected to the at least one propulsor (20), wherein preferably an additional electrical energy storage device (26) is arranged between the generator and the electric motor, which is designed for intermediate storage of the electrical energy generated via the generator. Modular engine (1) according to one of the preceding claims, wherein the at least one propulsor (20) in particular has a propeller. Modular engine (1) according to claim 5, wherein the propeller has foldable and extendable rotor blades. Modular engine (1) according to one of the preceding claims, additionally comprising an adjustment mechanism (18), wherein the adjustment mechanism (18) is configured to selectively direct the fluid flow from the outlet of the gas turbine (12) to the inlet of the payload turbine (14) or to the inlet of the nozzle assembly (16) for the activation or deactivation of the payload turbine (14). Modular engine (1) according to claim 7, wherein the adjustment mechanism (18) in particular comprises a valve or flap for redirecting the fluid flow exiting the outlet of the gas turbine (12). Modular engine (1) according to one of the preceding claims, wherein the nozzle assembly (16) has a variable surface geometry. Modular engine (1) according to one of the preceding claims, wherein the engine (1) is configured to activate the payload turbine (14) by redirecting the fluid flow and passing the fluid flow from the outlet of the gas turbine (12) through the payload turbine (14). Modular engine (1) according to one of the preceding claims, wherein the engine (1) is configured to deactivate the payload turbine (14) by diverting the fluid flow from the payload turbine (14) to the nozzle assembly (16) with the fluid flow from the outlet of the gas turbine (12), so that the payload turbine (14) is no longer subjected to the fluid flow. Modular engine (1) according to one of the preceding claims, wherein the engine (1) is configured to deactivate the payload turbine by jettisoning the drive train (22). Modular engine (1) according to one of the preceding claims, wherein the engine (1) additionally comprises a control unit, the control unit being configured to control the activation and deactivation of the payload turbine (14). Method for operating a modular engine (1), according to one of claims 1 to 13, with a first and at least one second drive phase of the engine (1), wherein in the first drive phase the payload turbine (14) is activated to drive the at least one propulsor (20) by passing the fluid flow from the outlet of the gas turbine (12) through the payload turbine (14) and in the at least second drive phase the payload turbine (14) is deactivated and the fluid flow from the outlet of the gas turbine (12) is expelled through the nozzle assembly (16), characterized in that the payload turbine (14) and / or the at least one propulsor (20) is jettisoned from the modular engine (1) when switching from the first drive phase to the second drive phase. Method according to claim 14, wherein in the first drive phase the thrust generation is primarily carried out by the at least one propulsor (20), wherein the at least one propulsor (20) is driven by the payload turbine (14) and in at least a second drive phase the thrust generation is primarily carried out by the emission of a fluid jet from the nozzle device (18). Method according to one of claims 14 or 15, wherein the operation of the engine (1) alternates at least once between the first and second drive phases. Use of the modular engine (1) according to one of claims 1 to 13 for carrying out the method according to one of claims 14 to 16.

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