LAUNCH ROCKET AND METHOD FOR OPERATING A LAUNCH ROCKET

Description

TECHNICAL FIELD

[0001] The present invention relates to a launch vehicle for space flights according to the preamble of patent claim 1. It further relates to methods for operating a launch vehicle. BACKGROUND OF THE INVENTION

[0002] Rockets intended for flight into space are usually designed with multiple stages, with a first propulsion stage equipped with recoil engines propelling the rocket into the upper regions of the troposphere, up into the upper region of the stratosphere or even beyond, to an altitude of around 50 to 70 km, where a second propulsion stage is then ignited, transporting the rocket into orbit or onto an interplanetary trajectory. The first propulsion stage falls back to Earth after the engines have burned out, and the second propulsion stage usually burns up upon re-entry into the atmosphere. For several years now, successful attempts have been made to land first propulsion stages equipped with recoil engines in a controlled manner after the end of their burnout so that they can be reused. Such reuse of rocket stages is desirable for economic reasons. STATE OF THE ART

[0003] CA 2 428 883 A1 shows and describes a single-stage space launch vehicle with a substantially square footprint, equipped with rotors at the four corners of its upper surface, each with two counter-rotating propellers. In addition, the launch vehicle is equipped with a central rocket motor for launch.

[0004] DE 4215835 A1 shows and describes a reusable single-stage spacecraft with a circumferential rotor arranged along the equator of the rotationally elliptical device body and a centrally provided and vertically downward-directed recoil drive as well as an obliquely downward-directed take-off recoil drive which accelerates the spacecraft during take-off after take-off by means of the circumferential rotor into an obliquely upward but predominantly horizontally directed trajectory in order to achieve aerodynamic lift forces on the device body.

[0005] US 2020 / 0262590 A1 shows and describes a vertically launched rocket with a cylindrical rocket stage and a separate launch stage, whose launch stage engines are arranged around the cylindrical rocket stage and are designed as air-breathing combustion engines.

[0006] DE 102014010109 A1 shows and describes a single-stage military missile which is not designed for space flight but for the transport of ballistic missiles and which has a cylindrical fuselage around which four engines equipped with electrically driven propellers are arranged.

[0007] US 2013 / 0087659 A1 shows and describes a space rocket having a first propulsion stage equipped with two wings projecting radially from the cylindrical fuselage, forming a tail unit and equipped with movable control fins at their radial outer ends. This allows this first rocket stage, after separation from the further-flying second rocket stage, to land on Earth using a landing gear like a winged aircraft. Electrically driven propellers can be provided for the return flight and landing.

[0008] WO 2017 / 021 758 A1 shows and describes a launcher for a high-altitude rocket launch. It comprises four radial structures arranged crosswise extending from the rocket, each of which has a helicopter-like, electrically driven rotor at its free end. The rotors' rotation axes run parallel to the rocket's longitudinal axis. The electric drives are powered externally via a cable leading from the ground to the airborne launcher.

[0009] US 2019 / 0 225 330 A1 shows and describes an unmanned aerial vehicle that, emanating from a central fuselage, has four crosswise arranged radial structures, each with a helicopter-like, electrically driven rotor at its free end. The electrical energy for this is stored in corresponding energy storage units in the aircraft. The rotors can be operated in a propulsion mode, but can also be operated in a generator mode for energy recovery.

[0010] DE 10 2015 014 502 A1 shows and describes an unmanned aerial vehicle with a fuselage equipped with a solid-fuel propulsion system and with an auxiliary wing system comprising at least one aerodynamic lift-generating wing. Electric motor drive systems equipped with a pusher propeller as propulsion propeller are provided on the wings.

[0011] DE 10 2014 019 398 A1 discloses a launch frame platform for a vertically launched rocket. The platform surrounds the rocket and has a plurality of rotor-like propulsion devices, the rotational axes of the rotors being aligned parallel to the longitudinal axis of the rocket. The rotor-like propulsion devices are equipped with conventional combustion engines. The launch frame platform transports the rocket to a level above the ground, from where the rocket is then launched by igniting the rocket engines. The launch frame platform then floats back to the Earth's surface on parachutes. PRESENTATION OF THE INVENTION

[0012] The object of the present invention is to provide an improved launch vehicle for flights into space which has a higher reuse rate.

[0013] This object is achieved by a launch vehicle having the features of claim 1.

[0014] A launch vehicle with a rocket body having a longitudinal axis and at least one propulsion stage driven by a recoil drive acting predominantly parallel to the longitudinal axis, wherein the launch vehicle is provided with a plurality of outer rotors each driven by a rotor drive, the respective rotor axis of which is aligned substantially parallel to the longitudinal axis of the rocket body, wherein a separate launch stage is provided which is coupled to the rocket body and / or the propulsion stage or can be coupled and decoupled therefrom, and which has the plurality of outer rotors, wherein the outer rotors are arranged radially outside the rocket body in the manner of a multicopter, surrounding it, and wherein the respective rotor drive has at least one electric motor as a drive machine, is characterized according to the invention in that the propulsion stage forms a main propulsion stage,which is coupled or can be coupled to and decoupled from the rocket body and / or to an upper stage of the launch vehicle having at least one recoil drive, that the main propulsion stage has at least one recoil drive and a plurality of inner rotors which can each be driven by a rotor drive, the respective rotor axes of which are aligned substantially parallel to the longitudinal axis of the rocket body and which are arranged in a ring around the longitudinal axis in the manner of a multicopter, and that the launch stage is provided with at least one power storage device for storing electrical energy and for supplying the outer rotor drives with electrical energy, and that at least some of the rotor drives can be operated in a generator mode in which electrical energy can be generated during autorotation of the outer rotors, and that the respective rotor drives are designed,to return the generated electrical energy to the respective assigned power storage device. ADVANTAGES

[0015] The inventive provision of a launch stage with electrically powered rotor drives makes it possible to initially transport the launch vehicle with its payload to a predetermined first altitude without the use of propulsion engines designed as rocket motors, only at this point the ignition of such propulsion engines of the actual propulsion stage takes place. The electrically powered rotors reduce noise during launch and do not generate engine exhaust gases near the ground. This reduces rocket fuel consumption and the amount of exhaust gases generated near the ground, as well as noise and exhaust emissions near the ground. When the launch stage returns to the Earth's surface, the launch stage's rotor drives can be used to actively approach a selected landing site for the launch stage and perform a controlled landing there.The ring-shaped arrangement of the outer rotors, similar to a multicopter, ensures effective, agile, and stable control of the launch vehicle during liftoff from the launch platform until it reaches its first altitude. This first altitude lies within a range where the atmospheric density is still high enough for vertical propulsion by the rotors to be effective enough for the launch vehicle connected to the launch stage to at least maintain the reached altitude. As the launch stage, detached from the launch vehicle, falls back to Earth, the rotors of the rotor drives are driven by the airflow created during the fall, causing the electric motor of the associated rotor drive to rotate. The rotor drives, which thus act as generators, can thus supply electrical energy to their associated storage device and charge it.This stored energy can then be used as propulsion energy for the rotor drives during the approach phase to the landing site and during landing. The launch stage, equipped with at least one power storage device, is energy-autonomous, so that the rotor drives can continue to operate even after the launch stage has been decoupled from the propulsion stage or the rocket body during the landing phase. The rotor drives provided in the main propulsion stage can support the rotor drives of the launch stage during launch and later, upon the return of the main propulsion stage, also enable a targeted approach to a landing site for the main propulsion stage and a controlled landing of the main propulsion stage, thus making the main propulsion stage reusable.The ring-shaped, multicopter-like arrangement of the inner rotors also ensures effective, agile and stable control of the launch vehicle during takeoff from the launch platform until it reaches the first altitude level.

[0016] Further preferred and advantageous design features of the launch vehicle according to the invention are the subject of subclaims 2 to 6.

[0017] According to an advantageous embodiment of the invention, the upper stage, which is equipped with at least one propulsion drive, also has a plurality of landing rotors, each driven by a rotor drive, whose respective rotor axes are aligned substantially parallel to the longitudinal axis of the rocket body and which are arranged in a ring inside the upper stage in the manner of a multicopter. The provision of these landing rotors also makes it possible to intercept the upper stage in a controlled manner after a controlled re-entry into the Earth's atmosphere and a subsequent fall to Earth using the then-driven landing rotors, and to fly it to a landing site and land there. This also makes the upper stage reusable.

[0018] It is particularly advantageous if the respective rotor drive incorporates at least one electric motor as the prime mover. Such electrically driven rotors reduce noise during takeoff and do not generate engine exhaust near the ground.

[0019] It is also advantageous if the main propulsion stage is also equipped with a power storage device for storing electrical energy and supplying electrical power to the inner rotors. This makes the main propulsion stage energy autonomous, and its rotor drives can also be operated after the launch stage has been decoupled from the propulsion stage or the rocket body during the landing phase.

[0020] It is also advantageous if the upper stage is equipped with a power storage device for storing electrical energy and supplying the inner rotors with electrical energy.

[0021] A particularly advantageous development of the launch stage and / or the main propulsion stage is characterized in that at least some of the respective rotor drives can be operated in a generator mode during autorotation of the associated rotors, in which electrical energy can be generated, and in that the respective rotor drives are designed to feed the generated electrical energy back into the respectively associated power storage device. When the respective stage, decoupled from the launch vehicle, falls back to Earth, the rotors of the rotor drives are driven by the air flow created during the fall and cause the electric motor of the respectively associated rotor drive to rotate. The rotor drives, thus functioning as generators, can thus supply electrical energy to their associated storage device and charge it.This stored energy can then be used as propulsion energy for the rotor drives during the approach phase to the landing site and during landing.

[0022] The part of the problem directed to the method for operating a launch vehicle is alternatively solved by the method according to claim 7 and the method according to claim 8.

[0023] In a first method for operating a launch vehicle with a reusable launch stage, the launch vehicle is transferred from a launch site to a predetermined first altitude using the launch stage's rotor drives. Upon reaching the first altitude, the propulsion stage's recoil drive is ignited, and the launch stage is decoupled from the rocket body or the propulsion stage. The launch stage then descends from the first altitude, recovering electrical energy through the rotor drives acting as generators, and is returned to Earth and a landing site.

[0024] In a second method for operating a launch vehicle equipped with a reusable launch stage and a main propulsion stage, the launch vehicle is first transferred from a launch site to a predetermined first altitude level using the rotor drives of the launch stage and the main propulsion stage. Upon reaching the first altitude level, the launch stage is decoupled from the rocket body or the main propulsion stage, respectively, and the recoil drive of the main propulsion stage is ignited. The launch stage then descends from the first altitude level—as in the first method—and is returned to Earth at a landing site. The launch vehicle then continues its flight using the recoil drive of the main propulsion stage, with the rotor drives of the main propulsion stage no longer contributing to propulsion and preferably being covered for their protection.If the launch vehicle subsequently reaches a second, higher altitude, the main propulsion stage is decoupled from the rocket body or the upper stage, and the upper stage's recoil engine is ignited. The main propulsion stage then descends from the second altitude and is returned to Earth and a landing site. If the second altitude is in space outside the Earth's atmosphere, the main propulsion stage initially re-enters the Earth's atmosphere. To protect against the heat generated during re-entry, the main propulsion stage is, in this case, at least partially covered with a heat shield or at least a thermal protection layer.

[0025] In a third method for operating a launch vehicle with a reusable launch stage, a main propulsion stage, and an upper stage, the launch vehicle is transferred from a launch site to a predetermined first altitude level by means of the rotor drives of the launch stage and the main propulsion stage. Upon reaching the first altitude level, the recoil drive of the main propulsion stage is ignited and the launch stage is decoupled from the rocket body or the main propulsion stage. The launch stage then descends from the first altitude level and is returned to Earth. When the launch vehicle reaches a second, higher altitude level, the main propulsion stage is decoupled from the rocket body or the upper stage, and the recoil drive of the upper stage is ignited. The main propulsion stage then descends from the second altitude level to Earth and is returned.The upper stage, which initially continues to fly, is transported back to Earth by means of at least one recoil propulsion device after reaching a target orbit and is returned to the Earth's surface after re-entering the Earth's atmosphere.

[0026] All three methods according to the invention enable a controlled landing and reusability of the launch stage, the second and third methods also enable a controlled landing and reusability of the main propulsion stage, and the third method finally enables the reuse of the upper stage.

[0027] In all three methods, it is advantageous if the rotors of the launch stage are operated in an autogyro mode during the descent within a region of the atmosphere where the air density is sufficiently high for rotor operation down to a first intercept altitude level, with at least part of the rotor devices generating electrical energy and feeding it back into the associated power storage device, which is thereby charged. Upon reaching the first intercept altitude level, the rotor devices of the launch stage are switched back to a propulsion mode, whereupon the launch stage is operated in a controlled descent and landing mode. The autogyro mode not only stabilizes the uncontrolled descent of the launch stage from the first altitude level to the first intercept altitude level, but also supplies electrical energy that can be used in the subsequent controlled descent, approach, and landing.

[0028] In the second method, it is advantageous if, in an analogous manner, the rotors of the main propulsion stage are operated in an autogyro mode during descent (within the atmosphere at sufficiently high air density) to a second interception altitude level, wherein at least part of the rotor devices generates electrical energy and feeds it back into the associated power storage device, and if, upon reaching the second interception altitude level, the rotor devices of the main propulsion stage are switched back to a propulsion mode, whereupon the main propulsion stage is operated in a controlled descent and landing mode.Here, too, the autogyro mode not only stabilizes the uncontrolled descent of the main propulsion stage from the second altitude level to the second intercept altitude level, but also supplies electrical energy to the associated power storage device, which can be used again to drive the rotors during the subsequent controlled descent, the approach to landing, and the landing.

[0029] Finally, in a third method, it is advantageous if the landing rotors of the upper stage, which are covered and protected by corresponding flaps during the journey into orbit and back into the atmosphere, are operated in an autogyro mode during the descent within the Earth's atmosphere to a third interception altitude level, wherein at least part of the rotor devices generates electrical energy and feeds it back into the associated power storage device, and if, upon reaching the third interception altitude level, the rotor devices of the upper stage are switched to a drive mode, whereupon the upper stage is operated in a controlled descent and landing mode.

[0030] The idea of the present invention is therefore primarily to provide, in the case of a vertically or essentially vertically launching launch vehicle, a separate launch stage in the manner of a multicopter, preferably with a battery-electric rotor or propeller drive, which carries the actual launch vehicle into the region of the lower stratosphere, where only then are the recoil drives, for example the rocket motors, ignited, whereupon the launch stage, detached from the launch vehicle, falls back to Earth and is returned in a controlled manner to the launch site or to another landing site by means of its rotor drives.During this fall back to Earth, the rotor drives are preferably operated in autorotation mode (autogyro mode), whereby the rotors or propellers are set in rotation by the air flowing towards them during the fall and whereby the rotor shaft drives a generator, preferably the electric drive motor acting as a generator, and thus generates electrical energy with which the power storage devices are charged for the driven operating state of the rotor drives required during the landing phase.

[0031] During the transfer to the stratosphere via the launch stage, additional, preferably battery-powered, rotor drives with associated rotors, which are provided in the first rocket stage, the main propulsion stage, can optionally provide support. Such rotor drives, provided in a rocket stage (main propulsion stage or upper stage), which can be constructed and function in the same way as those of the launch stage, can decelerate the rocket stage as it falls back to Earth with their respective rotors, preferably generating and storing electrical energy, and return it to the launch site or another landing site in a controlled landing flight with the actively driven rotor drives, thus making the rocket stage reusable.

[0032] Instead of batteries or accumulators, other power storage devices such as supercapacitors or other electrical energy sources such as fuel cells can of course also be used to supply energy to the rotor drives.

[0033] Preferred embodiments of the invention with additional design details and further advantages are described and explained in more detail below with reference to the accompanying drawings. BRIEF DESCRIPTION OF THE DRAWINGS

[0034] It shows: Fig. 1 is a vertical section through a first embodiment of a launch vehicle according to the invention; Fig. 2 is a horizontal section through a main propulsion stage of the first embodiment according to the invention along the line II-II in Fig. 1; Fig. 3 is a plan view of a launch stage designed according to the invention; Fig. 4 is a vertical section through a second embodiment of a launch vehicle according to the invention; Fig. 5 is a horizontal section through a main propulsion stage of the second embodiment designed according to the invention along the line VV in Fig. 4 and Fig. 6 a horizontal section through an upper stage of the second embodiment according to the invention along the line VI-VI in Fig. 4 . PRESENTATION OF PREFERRED EMBODIMENTS

[0035] Fig. 1 shows a vertical section through a first embodiment of a launch vehicle 1 according to the invention. The launch vehicle 1 comprises a rocket body 2 which has a main propulsion stage 3 and an upper stage 4 and which is decoupleably connected to a launch stage 5 in the region of its main propulsion stage 3.

[0036] The upper stage 4 is essentially formed by a cylindrical casing 40, which is provided with a foldable conical upper tip 41. In the upper region of the upper stage 4, a payload compartment 42 is formed for accommodating a payload 6, which is accessible by folding open the conical tip 41, so that the payload 6 can be released into space from the payload compartment 42.

[0037] In the lower region of the upper stage 4, i.e., on the underside facing away from the conical tip 41, a propulsion engine 43 is provided, the outlet nozzle 44 of which is directed downwards and which is arranged coaxially to the vertical longitudinal axis Z of the launch vehicle 1. Between the propulsion engine 43 and the payload compartment 42, a supply compartment 45 is provided, in which a plurality of propellant tanks 46, 46' are arranged, which contain the propellants for operating the propulsion engine 43 and which are connected to the propulsion engine 43 via corresponding propellant lines (not shown).

[0038] The lower portion 40' of the cylindrical housing shell 40 of the upper stage 4, facing away from the conical tip 41, engages in a matching cylindrical receiving opening 31 in the upper side of the housing shell 30 of the main drive stage 3 and is removably inserted there. The upper stage 4 is thus detachably connected to the main drive stage 3.

[0039] The casing shell 30 of the main propulsion stage 3 has a spherical sector-like shape with a convex lower wall 30' facing away from the upper stage 4. The reusable first rocket stage of the launch vehicle 1, formed by the main propulsion stage 3, has the shape of a flat truncated cone with a convex base, similar to an Apollo capsule. The outer diameter of the main propulsion stage 3 is significantly larger than the outer diameter of the cylindrical upper stage 4. In the example shown, the outer diameter of the main propulsion stage 3 is approximately four times the outer diameter of the upper stage 4.

[0040] In its radially outer region, near the largest circumferential edge 30" of the casing shell 30 of the main propulsion stage 3, vertical air ducts 32 are provided, running axially parallel to the longitudinal axis Z of the launch vehicle 1 and distributed over the circumference of the main propulsion stage 3. Preferably, eight vertical air ducts 32 are provided ( Fig. 2). In each of the air ducts 32, a propeller-like inner rotor 34 is arranged, driven by an electric rotor drive 33, whose respective inner rotor axis Z RI runs parallel to the longitudinal axis Z of the launch vehicle 1. The upper openings 32' and the lower openings 32" of the air ducts 32 can be closed in the region of the housing shell 30 by means of protective flaps (not shown).

[0041] Centrally located below the receiving opening 31 for the upper stage 4, the main propulsion stage 3 contains a central engine compartment 35 with a plurality of thrusters 36, each of whose outlet nozzles 36' opens downwards away from the payload compartment 42. The engine compartment 35, which is open downwards during operation of the thrusters 36, can be closed by protective flaps (not shown). These protective flaps close the engine compartment 35, particularly during a dive of the reusable main propulsion stage 3 back to Earth.

[0042] In an annular interior area 37 between the vertical air ducts 32 and the central engine compartment 35, fuel tanks 38, 38' for supplying the thrust drives 36 are arranged distributed around the circumference. Radially outside this annular interior area 37, electrical power storage devices 39 are provided for supplying energy to the electric rotor drives 33 in the circumferential direction between the air ducts 32.

[0043] Alternatively, the rotor drives 33 can also be supplied with electrical energy from power storage devices provided in the launch structure 5.

[0044] The main propulsion stage 3 rests with its convex lower wall 30' on a ring structure element 50 of the Fig. 3The drive stage 5 is shown in isolation and is decoupled there from the drive stage 5. This decoupled connection can be formed, for example, by means of retaining clips (not shown). These retaining clips can also be used to transfer electrical energy from the drive stage 5 to the main drive stage 3.

[0045] The launch stage 5 has a plurality of propeller-like outer rotors 52 arranged around the ring structure element 50, each of which is drivable by an electric rotor drive 53 for rotation about a respective outer rotor axis Z RA , which runs parallel to the longitudinal axis Z of the launch vehicle 1. The respective electric rotor drive 53 and the outer rotor 52 associated with it together form a respective rotor drive nacelle 54. Each rotor drive nacelle 54 is mechanically connected to the ring structure element 50 via an associated support structure 55.

[0046] As in Fig. 3As can be seen, in the example shown there, eight rotor drive nacelles 54 are arranged uniformly in a ring around the ring structure element 50 at a distance of 45° from one another, so that the launch stage 5 is designed in the manner of a multicopter, similar to an octocopter. On the underside of the ring structure element 50 facing away from the rocket body 30, a plurality of landing legs 56 are provided. Electrical power storage devices 58 are provided inside the ring structure element 50, which are or can be electrically connected to the rotor drives 53 via electrical lines (not shown).

[0047] In an alternative embodiment, the upper stage 4 can also be completely enclosed by the casing 30 of the main propulsion stage 3, which then has the cylindrical upper casing with the hinged tip 41. In this embodiment, the upper stage can be deployed together with the payload 6 provided in or on the upper stage 4 through the hinged tip 41.

[0048] The Fig. 1 to 3 The embodiment of the launch vehicle 1 shown forms a partially reusable launch vehicle system for small and medium payloads which are to be transported, for example, into a sun-synchronous orbit.

[0049] The flight schedule of the Fig. 1 to 3The first embodiment of the launch vehicle 1 shown is as follows: During launch, the rotors of the launch stage 5 and the main propulsion stage 3 are initially operated together, with the rotors 52 of the main propulsion stage 3 being powered by the power storage devices 58 of the launch stage 5 or from on-board power storage devices 39. The launch stage with the attached launch vehicle 1 ascends vertically upwards at a speed of, for example, 50 m / s (180 km / h) into the lower region of the stratosphere, for example to an altitude of approximately 15 km. At this altitude, the rocket engine, i.e. the recoil drive 43, of the main propulsion stage 3 is ignited and the retaining clamps between the launch stage 5 and the main propulsion stage 3 are released. At the same time, the power supply to the rotors of the main propulsion stage is switched to the internal power storage devices 39.The power storage devices 58 of launch stage 5 are now empty, and the rotor drives 53 of launch stage 5 are shut down, whereupon launch stage 5 enters a free fall. During this dive, the outer rotors 52 are set in rotation by the oncoming air (autorotation), which, on the one hand, slows the launch stage's fall and, by switching the rotor drives 53, formed by electric motors, to act as generators, partially recharges the power storage devices 58 of launch stage 5. At a suitable altitude, the rotor drives 53 are switched on again, and the dive is stopped. Launch stage 5 then lands in a controlled manner at the launch site or at another specified landing site.

[0050] Meanwhile, the main propulsion stage 3 continues its flight, with the rotors 34 and the rotor thrusters 33 of the main propulsion stage 3 still running as long as they generate thrust. After the thrusters 36 of the main propulsion stage 3 have burned out, the nose cone 41 is opened and the upper stage 4 with the payload 6 is deployed, continuing the flight. The upper stage 4 of this embodiment is not reusable.

[0051] The return of the also reusable main propulsion stage 3 can be achieved in two ways. Either the thrusters 36 of the main propulsion stage 3 are fired again, counter to the direction of flight, to perform a "boost-back" maneuver, and the main propulsion stage 3 then falls back to the launch site on a parabolic trajectory, or the main propulsion stage 3 falls back to Earth without a "boost-back" but then lands away from the original launch site, for example, on a ship. In both cases, after the last thruster firing of the thrusters 36, the protective flaps of the engine compartment 35 are closed, and the main propulsion stage 3 returns to the ground in free fall. As with launch stage 5, the fall is slowed by autorotation of the inner rotors 34 and the rotor drives 33, and the batteries of the main propulsion stage 3 are recharged.The dive is recovered by engaging the rotor engines 33, but unlike the launch stage, the landing takes place on a prepared air cushion. After the flight and return of the launch stage 5 and the main propulsion stage 3, they are checked, repaired if necessary, and prepared for the next flight.

[0052] The launch vehicle 1 according to the invention can also be used as a sounding rocket by installing a corresponding experiment package instead of the upper stage.

[0053] Fig. 4shows a second embodiment of a launch vehicle 101 according to the invention with a main propulsion stage 103 and an upper stage 104 carrying a payload 106. The unit comprising the main propulsion stage 103 and the upper stage 104 forming the rocket body 102 has a spherical sector-like shape and forms the shape of a flat truncated cone with a convex base, similar to an Apollo capsule. This launch vehicle 101 is also—as in the first embodiment—provided with a launch stage 105, which, like the one shown in Fig. 3 shown starting stage 5 of the first embodiment is designed; for the starting stage 105, therefore, the same statements apply as in connection with Fig. 3 described for launch stage 5.

[0054] The main drive stage 103 is also constructed analogously to the main drive stage 3 of the first embodiment, but the external shape of the main drive stage 103 differs from the first embodiment, since the housing shell 130 of the main drive stage 103 describes a hollow spherical sector with a convex outwardly curved lower wall 130' and a concave inwardly curved upper wall 130".

[0055] As in the first embodiment, the main propulsion stage 103 is provided with air ducts 132 running vertically in the lower region and axially parallel to the vertical longitudinal axis Z' of the launch vehicle 101, distributed over the circumference of the main propulsion stage 103. Preferably, eight vertical air ducts 132 are provided ( Fig. 5). In this second embodiment, the respective air duct 132 also runs from a respective lower opening in the convexly outwardly curved lower wall 130' upwards to an opening in the conical side surface 130‴ of the housing shell 130. Due to the two-stage design shown in this exemplary second embodiment, the upper part of the respective air duct 132 is curved obliquely outwards. The central engine compartment 135 with the recoil drives 136 and the annular interior region 137 with the fuel tanks 138, 138' contained therein correspond in their construction to the first embodiment, as shown in Fig. 5 can be seen.

[0056] In each of the air ducts 132, a propeller-like inner rotor 134 is arranged, driven by an electric rotor drive 133, whose respective inner rotor axis Z' RI runs parallel to the longitudinal axis Z' of the launch vehicle 101. The upper openings 132' and the lower openings 132" of the air ducts 132 can be closed in the region of the side surfaces 130‴ and the convex lower wall 130' by means of protective flaps (not shown).

[0057] In the second embodiment, electrical energy storage devices 139 are also provided radially outside this annular interior region 137 for supplying energy to the electrical rotor drives 133 in the circumferential direction between the air ducts 132. Alternatively, the rotor drives 133 can also be supplied with electrical energy by the energy storage devices 158 provided in the ring structure element 150 of the starting stage 105.

[0058] As in the first embodiment, the main drive stage 103 rests with its convex lower wall 130' on a ring structure element 150 of the starting stage 105 and is decoupleably connected to the starting stage 105 there. This decoupleable connection can be formed, for example, by means of retaining clips (not shown). These retaining clips can also be used to transmit electrical energy from the starting stage 105 to the main drive stage 103.

[0059] The upper stage 104 has a spherical sector-like shape similar to an Apollo capsule and is shaped like a flat truncated cone with a convex base formed by a convex lower wall 140', the curvature of which corresponds to the curvature of the concavely inwardly curved upper wall 130" of the main propulsion stage 103. The convex lower wall 140' of the upper stage 104 rests on the concavely inwardly curved upper wall 130" of the main propulsion stage 103. Decoupleable coupling means hold the upper stage 104 firmly to the main propulsion stage. Preferably, the convex lower wall 140' of the upper stage 104 is provided on the outside with a heat shield (not shown) that protects the upper stage 104 during re-entry into the Earth's atmosphere.

[0060] As with the first embodiment, this upper stage 104 is also provided with a foldable conical upper tip 141. In the upper region of the upper stage 104, a payload compartment 142 is formed for receiving the payload 106, which is accessible by folding open the conical tip 141, so that the payload 106 can be released into space from the payload compartment 142. The lateral wall 140" of the casing 140 of the upper stage 104 has a conical side surface that merges into the conical side surface of the main propulsion stage 103.

[0061] As shown in the horizontal section of the Fig. 6As can be seen, the upper stage 104 is also equipped in its radially outer region with essentially vertically extending air ducts 147, in each of which a landing rotor 149 is arranged, which can be driven by an electric rotor drive 148 and is rotatable about a respective rotor axis Z' L, which runs parallel to the longitudinal axis Z'. The upper openings and the lower openings of the air ducts 147 can be closed in the region of the lateral wall 140" of the housing shell 140 and the convex lower wall 140' of the upper stage 104 by means of protective flaps (not shown). The electric rotor drives 148 are electrically connected to electrical power storage devices 144 in the upper stage 104.

[0062] Around the payload compartment 142, as shown in Fig. 6As can be seen, four thrusters 143 are provided coaxially to the vertical longitudinal axis Z of the launch vehicle 1, each of whose outlet nozzles is directed downwards. The outlet openings (not shown) surrounding the respective outlet nozzle in the convex lower wall 140' of the upper stage 104 can be closed by protective flaps (not shown).

[0063] Between the recoil drives 143 and the paired air ducts 147, fuel tanks 146, 146' are arranged, which contain the fuel for operating the recoil drives 143 and which are connected to the recoil drives 143 via corresponding fuel lines (not shown).

[0064] The Fig. 4 to 6 The illustrated embodiment of the launch vehicle 101 forms a fully reusable launch vehicle system for small and medium payloads, which are to be transported, for example, into a sun-synchronous orbit.

[0065] The flight schedule of the Fig. 4 to 6 The second embodiment of the launch vehicle 101 shown is as follows: The flight sequence of the launch stage 105 and the main propulsion stage 103 corresponds to the sequence described in connection with the first embodiment, wherein the protective flaps for the air ducts 147 are closed.

[0066] The upper stage 104 with the payload 106 reaches orbital velocity. After the payload 106 is deployed into orbit, the upper stage 104 is decelerated by re-igniting its thrusters 143 until it re-enters the atmosphere in a ballistic trajectory. Before re-entry, the protective flaps for the thrusters 143 are closed. The protective flaps for the air ducts 147 remain closed. The landing of the upper stage 104 occurs, as with the main propulsion stage 103, by opening the protective flaps for the air ducts and activating the rotor drives 148. The landing is then reversed using the rotor drives 148, as described analogously in connection with the main propulsion stage. The controlled landing also occurs on a prepared air cushion.

[0067] The launch vehicle 101 of the second embodiment can also be used for returnable payloads that are permanently installed in the upper stage 104 and return to Earth together with the upper stage 104 after the end of the mission.

[0068] Reference symbols in the claims, the description and the drawings serve only to improve the understanding of the invention and are not intended to limit the scope of protection. List of reference symbols

[0069] They refer to: 1 Launch vehicle 2 Rocket body 3 Main propulsion stage 4 Upper stage 5 Launch stage 30 Casing shell of the main propulsion stage 3 30' Convex lower wall 30" Circumferential edge of the casing shell 30 31 Cylindrical receiving opening 32 Air ducts 32' Upper openings 32" Lower openings 33 Electric rotor drive 34 Inner rotor 35 Central engine compartment 36 Recoil drive 37 Annular interior area 38 Fuel tank 38' Fuel tank 39 Electric power storage device 40 Cylindrical casing shell 40' Lower area of the cylindrical casing shell 40 41 Conical tip 42 Payload compartment 43 Recoil drive 44 Exit nozzle 45 Service compartment 46 Fuel tank 46' Fuel tank 50 Annular structural element 52 Outer rotor 53Electric rotor drive 54Rotor drive nacelle 55Support structure 56Landing leg 58Power storage devices 101Launcher 102Rocket body 103Main propulsion stage 104Upper stage 105Launch stage 106Payload 130'Lower wall 130"Upper wall 130"Side surface 132Air duct 132'Upper opening of the air duct 132132" lower opening of the air duct 132 133 electric rotor drive 134 inner rotor 135 central engine compartment 136 recoil drive 137 annular interior area 138 fuel tank 138' fuel tank 139 electric power storage device 140 upper stage casing 104 140' lower wall of the upper stage 104 140" side wall of the casing 140 141 conical tip 142 payload compartment 143 recoil drive 146 fuel tank 146' fuel tank 147 air duct 148 electric rotor drive 149 landing rotor 150 annular structural element 158 power storage devices ZLongitudinal axis Z'vertical longitudinal axis Z' L Rotor axis Z RI inner rotor axis Z' RI inner rotor axis Z RA outer rotor axis

Claims

1. A launch vehicle with a rocket body (2; 102) having a longitudinal axis (Z, Z'), which rocket body has at least one propulsion stage (3; 103) which can be driven by a reaction propulsion system (36; 136) acting predominantly parallel to the longitudinal axis (Z, Z'), wherein the launch vehicle is provided with a plurality of rotors each drivable by means of a rotor drive, the respective rotor axis of which is aligned substantially parallel to the longitudinal axis of the rocket body, wherein the launch vehicle comprises a separate takeoff stage (5; 105) coupled to or couplable to and decouplable from the rocket body (2; 102) and / or the propulsion stage (3; 103), said takeoff stage comprising the plurality of outer rotors (52), wherein the outer rotors (52) are arranged in the manner of a multicopter radially outside the rocket body (2; 102) surrounding it and wherein the respective rotor drive (53) has at least one electric drive motor as drive engine, wherein the propulsion stage forms a main propulsion stage (3; 103) which is coupled to or can be coupled to and decoupled from the rocket body (2; 102) and / or an upper stage (4; 104) of the launch vehicle (1; 101), which upper stage (4; 104) has at least one reaction propulsion system (43; 143), wherein the main propulsion stage (3; 103) has at least one reaction propulsion system (36; 136) and a plurality of inner rotors (34; 134) each of which can be driven by means of a rotor drive (33; 133) and whose respective rotor axis (ZRI, Z'RI) is aligned essentially parallel to the longitudinal axis (Z, Z') of the rocket body (2; 102), and which are arranged annularly around the longitudinal axis (Z, Z') in the manner of a multicopter, and wherein the takeoff stage (5; 105) is provided with at least one power storage device (58; 158) for storing electrical energy and for supplying electrical energy to the outer rotor drives (53), and wherein at least some of the rotor drives (53) can be operated in a generator mode, in which electrical energy can be generated, during an autorotation of the outer rotors (52), and wherein the respective rotor drives (53) are designed to feed the electrical energy generated back into the respectively associated power storage device (58; 158).

2. The launch vehicle according to claim 1, characterized in that the upper stage (104) is provided with at least one reaction propulsion system (143) and has a plurality of internal landing rotors (149) which can each be driven by means of a rotor drive (148) and whose respective rotor axis (Z'L) is aligned essentially parallel to the longitudinal axis (Z') of the rocket body (102), and which are arranged annularly around the longitudinal axis (Z, Z') in the manner of a multicopter in the interior of the upper stage.

3. A launch vehicle according to claim 1 or 2, characterized in that the respective rotor drive (33, 133, 148) has at least one electric drive motor as drive engine.

4. A launch vehicle according to claim 1 and 3 or any of claims 1 to 3, characterized in that the main propulsion stage (3; 103) is provided with at least one power storage device (39; 139) for storing electrical energy and for supplying the inner rotor drives (33; 133) with electrical energy.

5. A launch vehicle according to claim 2 and 3 or any of the claims 2 to 4, characterized in that the upper stage (104) is provided with at least one power storage device (144) for storing electrical energy and for supplying electrical energy to the rotor drives (148) of the inner landing rotors (149).

6. A launch vehicle according to any one of claims 4 or 5, characterized in that at least some of the rotor drives (33, 133; 148) can be operated in a generator mode, in which electrical energy can be generated, during an autorotation of the associated rotors (34; 134), and in that the respective rotor drives (33, 133; 148) are designed to return the electrical energy generated to the respectively associated power storage device (39, 139; 144).

7. A method of operating a launch vehicle (1; 101) according to claim 1 with a reusable takeoff stage (5; 105), wherein the launch vehicle (1; 101) is transferred in a first step by means of the rotor drives (53) of the takeoff stage (5; 105) from a launch site to a predetermined first altitude level, wherein on reaching the first altitude level the reaction propulsion system (36; 136) of the propulsion stage (3; 103) is ignited and the takeoff stage (5; 105) is decoupled from the rocket body (2; 102) or from the propulsion stage (3; 103), and the takeoff stage (5; 105) descends from the first altitude level with recovery of electrical energy by means of the rotors operating as generators and is returned again.

8. A method of operating a launch vehicle (1; 101) according to claim 1 with a reusable takeoff stage (5; 105) and a main propulsion stage (3; 103), wherein the launch vehicle (1; 101) is transferred in a first step by means of the rotor drives (53, 33, 133) of the takeoff stage (5; 105) and the main propulsion stage (3; 103) from a launch site to a predetermined first altitude level, the reaction propulsion system (36; 136) of the main propulsion stage (3; 103) being ignited and the takeoff stage (5; 105) being decoupled from the rocket body (2; 102) and from the main propulsion stage (3; 103) respectively when reaching the first altitude level, and the takeoff stage (5; 105) descending from the first altitude level and being returned again, wherein upon reaching a second altitude level, the reaction propulsion system (43; 143) of the upper stage (4; 104) is ignited and the main propulsion stage (3; 103) is decoupled from the upper stage (4; 104), and wherein the main propulsion stage (3; 103) descends from the second altitude level and is returned.

9. A method of operating a launch vehicle (101) according to claim 2 with a reusable takeoff stage (105),a main propulsion stage (103) and an upper stage (104), wherein the launch vehicle (101) is transferred in a first step by means of the rotor drives (53, 133) of the takeoff stage (105) and the main propulsion stage (103) from a launch site to a predetermined first altitude level, wherein, upon reaching the first altitude level, the reaction propulsion system (136) of the main propulsion stage (103) is ignited and the takeoff stage (105) is decoupled from the rocket body (101) and from the main propulsion stage (103), respectively, and wherein the takeoff stage (105) descends from the first altitude level and is returned, wherein upon reaching a second altitude level, the reaction propulsion system (143) of the upper stage (104) is ignited and the main propulsion stage (103) is decoupled from the upper stage (104), and wherein the main propulsion stage (103) descends from the second altitude level and is returned, and wherein the upper stage (104) is propelled back to earth by means of at least one reaction propulsion device after reaching a target orbit and is returned to the earth's surface after re-entering the earth's atmosphere.

10. A method according to claim 7 or 8, characterized in that the rotors (52) of the takeoff stage (5; 105) are operated in an autogyro mode when descending to a first intercept altitude level, at least a portion of the rotor drives (53) generating electrical energy and feeding it back into the associated power storage device (58; 158), and in that upon reaching the first intercept altitude level, the rotor drives (53) of the takeoff stage (5; 105) are switched back to a drive mode, whereupon the takeoff stage (5; 105) is operated in a controlled descent and landing mode.

11. A method according to claim 8, 9 or 10, characterized in that the rotors (34; 134) of the main propulsion stage (3; 103) are operated in an autogyro mode when descending to a second intercept height level, at least part of the rotor drives (33; 133) generating electrical energy and feeding it back into the associated power storage device (39; 139), and in that upon reaching the second intercept altitude level, the rotor drives (33; 133) of the main propulsion stage (3; 103) are switched back to a drive mode, whereupon the main propulsion stage (3; 103) is operated in a controlled descent and landing mode.

12. A method according to claim 9, 10 or 11, characterized in that the landing rotors (149) of the upper stage (104) are operated in an autogyro mode during descent within the earth's atmosphere to a third intercept altitude level, at least a portion of the rotor drives (148) generating and returning electrical energy to the associated power storage device (144), and in that upon reaching the third intercept altitude level, the rotor drives (148) of the upper stage (104) are switched to a drive mode, whereupon the upper stage (104) is operated in a controlled descent and landing mode.