Propulsion stage of a launch vehicle, launch vehicle and method for operating a launch vehicle

Abstract

The invention relates to a propulsion stage of a launch vehicle comprising a rocket body (2, fig. 2) having a longitudinal axis (Z, fig. 1), wherein the propulsion stage (3, fig.1) is provided with at least one jet propulsion drive (36; 36A, 36B) acting predominantly parallel to the longitudinal axis (Z), which can be supplied with fluid fuel from at least one fuel tank (38, 38') by means of at least one associated fuel pump (50, 52), and wherein the propulsion stage (3, fig. 1) has a plurality of rotor arrangements (33) which can each be driven by a rotor drive, which has at least one motor (33'; 33'A, 33'B), characterised in that the motor (33'; 33'A, 33'B) of at least one of the rotor drives is designed for optionally driving the rotor arrangement (33A, 33B) and / or a fuel pump (50, 52) of the at least one jet propulsion drive (36; 36A, 36B).

Description

[0001] Propulsion stage of a launch vehicle, launch vehicle and method for operating a launch vehicle

[0002] TECHNICAL FIELD

[0003] The present invention relates to a propulsion stage, in particular a reusable propulsion stage, of a launch vehicle according to the preamble of claim 1. It further relates to a launch vehicle having at least one such propulsion stage and a method for operating such a launch vehicle.

[0004] BACKGROUND OF THE INVENTION

[0005] Launch vehicles intended for flight into space are typically multi-stage, with a first propulsion stage equipped with recoil thrusters propelling the launch vehicle into the upper regions of the troposphere, up to the upper region of the stratosphere, or even beyond, at approximately 50 to 70 km altitude. A second propulsion stage is then ignited, transporting the launch vehicle into orbit or onto an interplanetary trajectory. The first propulsion stage falls back to Earth after the engines burn out, and the second propulsion stage typically burns up upon re-entry into the atmosphere.

[0006] For several years, successful attempts have been made to land first stages equipped with thrusters in a controlled manner after burnout, allowing them to be reused. Such reuse of rocket stages is desirable for economic reasons. The reusable rocket stages known to date land with the aid of controlled operation of their rocket engines and, to ensure sufficient deceleration of these thrusters intended for rocket launch, must be brought into a suitable position during the fall back to Earth. This requires the provision of additional control nozzles. Furthermore, fuel reserves must still be available for the planned landing upon return to Earth (and possibly even upon re-entry into the Earth's atmosphere).

[0007] STATE OF THE ART

[0008] WO 2020 / 094640 A1 discloses a two-stage launch vehicle for space flight. In addition to a thruster, it is equipped with electric ducted fans along the cylindrical first-stage fuselage. Their duct and rotation axes extend orthogonally to the longitudinal axis of the launch vehicle. These fans can act as generators during the stage's return to Earth, recharging the batteries for the fan motors. The fans facilitate a controlled landing of the returned stage. The provision of electric fans and the associated fan motors in addition to a thruster increases the mass to be transported into space or higher atmospheric layers, thus reducing the launch vehicle's payload.

[0009] RU 2 603 305 C1 and RU 2 609 539 C1 each show and describe a reusable rocket system for manned space flight. This rocket system has a plurality of thrusters arranged at the rear, lower end of the central tubular rocket fuselage. Four gas turbine engines are arranged around this rocket fuselage in the lower section, serving as jet engines. A fuel pump, driven by its own drive, supplies the combustion chamber of each gas turbine with fuel from a fuel tank. A different fuel pump supplies each thruster with fuel.

[0010] EP 2 415 966 ​​A1 relates to a drive train for a gas turbine, but not to a thruster designed as a rocket engine. SUMMARY OF THE INVENTION

[0011] The object of the present invention is to provide a launch vehicle propulsion stage of the type mentioned above, as well as a launch vehicle equipped therewith, which is mass-optimized and minimizes payload loss. Furthermore, the object of the invention is to provide a method for operating such a launch vehicle.

[0012] The part of the problem directed to the drive stage is solved by a drive stage having the features of patent claim 1.

[0013] A propulsion stage of a launch vehicle, in particular a reusable propulsion stage, with a rocket body having a longitudinal axis is provided with at least one recoil drive operating predominantly parallel to the longitudinal axis, to which liquid propellant can be supplied from at least one propellant tank by means of at least one associated propellant pump. The propulsion stage has a plurality of rotor assemblies, each of which can be driven by a rotor drive having at least one motor. This propulsion stage is characterized in that the motor of at least one of the rotor drives is designed to selectively drive the rotor assembly and / or a fuel pump of the at least one recoil drive.

[0014] A thruster is understood here to be a generally known arrangement comprising a combustion chamber and a rigid or pivoting thrust nozzle, and a rotor arrangement is understood to be an atmospheric propeller with propellers or rotor blades that can be driven by a rotor drive.

[0015] ADVANTAGES

[0016] The inventive dual use of the respective motor for driving the fuel pump and / or the rotor assembly gives the motor a dual function, making it possible to use only one drive motor for the fuel pump and the rotor drive, thereby saving mass and thus weight. It is particularly advantageous if the motor of the at least one rotor drive can be coupled either to a drive shaft of the associated rotor drive or to a drive shaft of a fuel pump of the at least one recoil drive.

[0017] If the propulsion stage comprises a plurality of recoil drives arranged around the longitudinal axis and a plurality of rotor assemblies arranged around the longitudinal axis, the latter being arranged in a ring around the longitudinal axis, for example, in the manner of a multicopter, it is advantageous if each recoil drive is assigned a rotor assembly, preferably radially outside thereof. It is particularly advantageous if each pair of recoil drives is assigned a pair of rotor assemblies. This allows, if both recoil drives of a pair are supplied with fuel components (fuel and oxidizer) from the same fuel tanks - one for the fuel and one for the oxidizer - the motor of one rotor assembly can drive the fuel pump for the fuel and the other motor can drive the fuel pump for the oxidizer, and can be coupled to them for this purpose.

[0018] The invention is not limited to rocket propellants with two liquid propellant components (diergols), but can also be used in an analogous manner with only one propellant component (monergols) or more than two propellant components (e.g. triergols), wherein each liquid propellant component is assigned at least one propellant pump.

[0019] Further preferred and advantageous design features of the drive stage according to the invention are the subject of subclaims 2 to 7.

[0020] In a preferred embodiment of the invention, the respective motor is assigned a first switchable clutch for coupling the motor to a drive shaft of the associated rotor assembly, and a second switchable clutch for coupling the motor to a drive shaft of the associated fuel pump. This simple mechanical coupling is reliable, and the switchable clutches can be controlled by an electronic control or regulating device.

[0021] According to a particularly preferred variant of the invention, which can be combined with other variants, the clutches assigned to the respective motors can be switched alternately. This makes it possible to operate the propulsion stage only with the recoil drive during the launch vehicle's ascent, and to operate only the rotor assemblies during the fall back to Earth and during the landing phase.

[0022] Preferably, the respective motor has an output shaft that is connected or connectable to the first switchable clutch in the region of its first end and to the second switchable clutch in the region of its second end. This creates a simple and reliable mechanical arrangement that enables both switching the motor output to selectively couple it to the fuel pump or to the rotor assembly, as well as jointly coupling the motor output to the fuel pump and the rotor assembly. Such dual operation can be carried out during the ascent phase of the propulsion stage and thus of the launch vehicle, particularly during the transition between a rotor-driven ascent phase and a recoil-driven ascent phase.

[0023] Preferably, the respective motor is formed by an electric motor, and at least one power storage device is provided for storing electrical energy and supplying the respective electric motor with electrical energy. Such electrical operation is locally emission-free and quieter than operation of an internal combustion engine or a gas turbine as a motor. The power storage device can comprise accumulators and / or capacitors, for example, so-called supercapacitors (supercapacitors), but it can also comprise at least one fuel cell for generating electrical energy.It is particularly advantageous if at least some of the electric motors, when coupled to an associated rotor drive and driven by an associated rotor arrangement, can be operated in a generator mode in which electrical energy can be generated, and if the respective electric motors are designed to feed the generated electrical energy back into the respectively associated power storage device. This embodiment of the invention, which can be combined with other embodiments, allows the power storage device(s) to be charged during the free fall of the drive stage in the atmosphere during the fall-back phase, wherein the rotors of the rotor arrangements are set into autorotation by the airflow and the associated rotor drive functions as a generator and generates electrical power, which is stored in the power storage devices.

[0024] A further embodiment of the invention, which can be combined with other embodiments, provides that the power of the respective motor can be controlled or regulated, in particular via a speed control, both when driving the associated rotor assembly and when driving the associated fuel pump, for which purpose a control or regulation device is provided. This makes it possible not only to control or regulate the thrust generated by the rotor assembly, but also to control or regulate the thrust of the associated recoil drive via the power of the fuel pump.In particular, if the propulsion stage comprises a plurality of thrusters arranged around the longitudinal axis and a plurality of rotor assemblies arranged around the longitudinal axis, the latter being arranged in a ring around the longitudinal axis, for example, in the manner of a multicopter, and if each rotor assembly and each fuel pump is assigned a common motor, the direction of movement of the propulsion stage can be controlled by controlling or regulating the individual motors, for example by pumping more fuel into the thrusters on one side of the propulsion stage than on the opposite side of the propulsion stage. This makes it possible, for example, to use simple thrusters with rigidly installed thrusters instead of more complex swiveling nozzles, thereby reducing the complexity and cost of the propulsion stage.

[0025] The invention also includes a launch vehicle having at least one propulsion stage, in one of the above embodiments or combinations thereof.

[0026] The part of the problem to be solved which is directed towards the method is solved by a method for operating a launch vehicle according to the invention, wherein the propulsion stage is driven by the at least one recoil drive in an ascent phase, wherein the propulsion stage, after the end of the ascent phase, falls back to Earth without propulsion in a fall phase, and wherein the propulsion stage, shortly before reaching the Earth's surface, is transferred into a landing phase in which it heads for a landing position in controlled flight and lands in a controlled manner, and in which at least some of the motors are coupled to a respectively assigned fuel pump in the ascent phase, and in which at least some of the motors are coupled to a respectively assigned rotor arrangement in the landing phase and drive the assigned rotor arrangement in a controlled manner.In this method according to the invention, the propulsion stage is intercepted at the end of the fall back to Earth phase by activating the motors coupled to the rotor assemblies by the lift force generated by the rotors of the rotor assemblies and the propulsion stage can then be flown in a precisely controlled manner to a landing site by means of the rotor assemblies and landed there.

[0027] A particularly advantageous development of the method according to the invention is one in which at least some of the motors designed as electric motors are coupled to a respective associated rotor arrangement during the return phase and are operated in a generator mode. The associated rotor arrangements drive the associated motor acting as a generator through autorotation of their rotors, and the electrical energy generated thereby is fed back into a respective associated power storage device and stored there. In this variant of the method according to the invention, electrical energy is generated when the drive stage returns to earth, which energy is used to charge the power storage devices.This allows the electrical energy consumed during the ascent phase, for example, to operate the fuel pumps, to be replaced by electrical energy recovered during the descent through the atmosphere. This allows the engines to have full power available during the landing phase, supplying the rotor assemblies with electrical energy during the return flight to the landing site and during landing. This increases the range for the return flight to the landing site.

[0028] 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.

[0029] BRIEF DESCRIPTION OF THE DRAWINGS

[0030] It shows:

[0031] Fig. 1 is a vertical section through a first embodiment of a drive stage according to the invention along the line 11 in Fig. 2;

[0032] Fig. 2 is a horizontal section through the drive stage designed according to the invention along the line ll-ll in Fig. 1 and

[0033] Fig. 3 is a schematic drawing of a drive group comprising two recoil drives and two rotor assemblies according to detail III in Fig. 2.

[0034] PRESENTATION OF PREFERRED EMBODIMENTS

[0035] Fig. 1 shows a vertical section through a launch vehicle 1 with a rocket body 2, which has a propulsion stage 3 and an upper stage 4. The upper stage 4 is essentially formed by a cylindrical casing 40, which is provided with a foldable or unfoldable 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.

[0036] 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 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 of the upper stage 4 and which are connected to the propulsion engine 43 of the upper stage 4 via corresponding propellant lines (not shown).

[0037] 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 drive stage 3 and is removably inserted there. The upper stage 4 is thus detachably connected to the drive stage 3.

[0038] The casing shell 30 of the 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 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 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 propulsion stage 3 is approximately four times as large as the outer diameter of the upper stage 4. In its radially outer region, near the largest circumferential edge of the casing shell 30 of the propulsion stage 3, a plurality of rotor assemblies 33 are provided, axially parallel to the longitudinal axis Z of the launch vehicle 1, spaced apart from one another in the circumferential direction, the respective rotor axes Z Rparallel to the longitudinal axis Z of the launch vehicle 1. The rotor assemblies 33 are located within the housing shell 30. The housing shell 30 has a plurality of vertically extending air ducts 32 distributed over the circumference of the propulsion stage 3, in each of which a rotor assembly 33 is arranged. 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).

[0039] Alternatively, the rotor assemblies 33 are covered by radially outwardly movable casing sections 30" (Fig. 2), which can be moved radially outward for the operation of the associated rotor assembly 33, for example on rails 30"', as shown for the left casing section 30"L with dashed lines in Fig. 2. In this alternative variant, a vertical air duct 32'" is also formed between each outwardly moved casing section 30" and the remaining central part of the housing casing 30, in which at least one respective rotor assembly 33 is located. In the example shown in Fig. 2, a pair of such rotor assemblies is located in an axially parallel air duct 32'" formed in this way, with four radially outwardly movable casing sections 30" being provided, distributed over the circumference at an angle of 90° to one another.

[0040] Preferably, eight propeller-like rotor assemblies 33 are provided, each electrically driven by a rotor drive with a motor 33' designed as an electric motor (Fig. 2). Each of these rotor assemblies 33 has an upper rotor 34 and a lower rotor 34', which, in an operating state driven by the motor 33', can be driven in the same direction or in opposite directions (depending on the type of propeller-like rotor 34, 34') to generate a vertical airflow - in the lift direction or in the downforce direction. Radially within the rotor assemblies 33, a plurality of thrust drives 36 are provided in a respective engine compartment 35, with each thrust drive 36 being assigned to a rotor assembly 33. The outlet nozzle 36' of the respective thrust drive 36, forming a thrust nozzle, is directed away from the payload compartment 42 and opens downwards.The engine compartment 35, which is open downwards during operation of the recoil propulsion units 36, can be closed by at least one protective flap (not shown). These protective flaps close the respective engine compartment 35, particularly during a dive of the reusable propulsion stage 3 back to Earth.

[0041] In a central interior area 37 radially inside and above the engine compartments 35, a central fuel tank 38 for storing a fuel and an annular fuel tank 38' for storing an oxidizer for supplying the thrusters 36 are arranged.

[0042] In Fig. 3, according to detail III in Fig. 2, a drive group 39 comprising two recoil drives 36A, 36B with a common fuel supply device 5 and two rotor assemblies 33A, 33B is schematically shown. Each of the rotor assemblies is assigned a motor 33'A, 33'B designed as an electric motor as the rotor drive. The output shaft 33"A, 33"B of a respective motor 33'A, 33'B can be coupled at one end via an associated first coupling 33"'A, 33"'B to a drive shaft 33""A, 33""B of the respective rotor assembly 33A, 33B for torque transmission.

[0043] The fuel supply device 5 has a first fuel pump 50 for the fuel and a second fuel pump 52 for the oxidizer. These fuel pumps 50, 52 are typically designed as turbopumps with a stator and a rotor rotating with a drive shaft 51, 53. The respective fuel pump 50, 52 is fluidly connected to the respective associated fuel tank 38, 38' via a feed line 54, 55 with a respective fluid inlet. From a respective fuel outlet of each fuel pump 50, 52, a first common fuel pressure line 56 for the fuel and a second common fuel pressure line 57 for the oxidizer lead to the two associated recoil drives 36A, 36B.The drive shaft 51 of the first fuel pump 50 for the fuel can be coupled via an associated coupling 58 to the other end of the output shaft 33"B of the motor 33'B for torque transmission, and the drive shaft 53 of the second fuel pump 52 for the oxidizer can be coupled via an associated coupling 59 to the other end of the output shaft 33'A of the motor 33'A for torque transmission.

[0044] The respective motor 33'A, 33'B can therefore be switched between a drive connection with the respectively assigned rotor arrangement 33A, 33B for driving the rotors 34, 34' and a drive connection with the respectively assigned fuel pump 50, 52 by means of a control or regulating device 60 that supplies control signals to the clutches 33"'A, 33"'B, 58, 59 and preferably also to the motors 33'A, 33'B. In special cases, the output shaft 33"A, 33"B of the respective motor 33'A, 33'B can also be coupled to both the assigned rotor arrangement 33A, 33B and the assigned fuel pump 50, 52 by closing both clutches 33"'A, 58 and 33"'B, 59 assigned to it.Even in the fall-back phase of the drive stage, the respective motor 33'A, 33'B is coupled by a closed coupling connection to the drive shaft 33""A, 33""B of the respectively associated rotor arrangement 33A, 33B for torque transmission, so that the rotors 34, 34' of the respective rotor arrangement 33A, 33B, driven for autorotation by the air flow impinging on the falling drive stage 3, drive the motor 33'A, 33'B, which now acts as a generator. The electrical current generated by the motor 33'A, 33'B is stored in at least one associated power storage device 62 (Fig. 2).

[0045] Reference symbols in the claims, the description, and the drawings serve only to facilitate understanding of the invention and are not intended to limit the scope of protection. List of reference symbols

[0046] They refer to:

[0047] 1 launch vehicle

[0048] 2 rocket bodies

[0049] 3 drive stage

[0050] 4 Upper School

[0051] 5 Fuel supply system

[0052] 6 Payload

[0053] 30 Housing shell

[0054] 30' convex lower wall

[0055] 30" radially outwardly movable jacket sections

[0056] 30"L left outwardly movable jacket section

[0057] 30" rails

[0058] 31 adapted cylindrical receiving opening

[0059] 32 air duct

[0060] 32' upper opening of the air duct

[0061] 32" lower air duct opening

[0062] 32'" air duct

[0063] 33 Rotor arrangement

[0064] 33A rotor arrangement

[0065] 33B Rotor arrangement

[0066] 33' engine

[0067] 33'A engine

[0068] 33' B Engine

[0069] 33"A Drive shaft of the engine 33'A

[0070] 33"B Drive shaft of the engine 33'B

[0071] 33"'A first coupling

[0072] 33"'B first coupling

[0073] 33""A Drive shaft of the rotor assembly 33A

[0074] 33""B Drive shaft of the rotor assembly 33B 34 upper rotor

[0075] 34' lower rotor

[0076] 35 Engine room

[0077] 36 recoil drive

[0078] 36A recoil drive

[0079] 36B recoil drive

[0080] 36' outlet nozzle

[0081] 37 central interior area

[0082] 38 central fuel tank

[0083] 38' annular fuel tank

[0084] 39 Drive group

[0085] 40 cylindrical housing shell

[0086] 40' lower section of the cylindrical casing shell

[0087] 41 conical upper tip

[0088] 42 payload space

[0089] 43 Recoil drive

[0090] 44 Outlet nozzle

[0091] 45 Supply room

[0092] 46 Fuel tank

[0093] 46' fuel tank

[0094] 50 first fuel pump

[0095] 51 Drive shaft

[0096] 52 second fuel pump

[0097] 53 Drive shaft

[0098] 54 feed line

[0099] 55 feed line

[0100] 56 first common fuel pressure line

[0101] 57 second common fuel pressure line

[0102] 58 Clutch

[0103] 59 Clutch

[0104] 60 Control or regulating device

[0105] 62 Power storage device Z vertical longitudinal axis

[0106] Z R Rotor axis

Claims

Claim 1. Propulsion stage of a launch vehicle, in particular a reusable propulsion stage, with a rocket body (2) having a longitudinal axis (Z), wherein the propulsion stage (3) is provided with at least one reaction drive (36; 36A, 36B) acting predominantly parallel to the longitudinal axis (Z), to which liquid propellant can be supplied from at least one propellant tank (38, 38') by means of at least one associated propellant pump (50, 52), and wherein the propulsion stage (3) has a plurality of rotor assemblies (33), each of which can be driven by a rotor drive comprising at least one motor (33'; 33'A, 33'B), characterized in that the motor (33'; 33'A, 33'B) can selectively drive at least one of the rotor drives for the rotor assembly (33A, 33B) and / or a propellant pump (50, 52) of the at least one reaction drive (36; 36A, 36B). 36B). 2.Drive stage according to claim 1, characterized in that each motor (33'; 33'A, 33'B) is assigned a first switchable clutch (33"'A, 33"'B) for coupling to a drive shaft (33"'A, 33"'B) of the associated rotor assembly (33) and a second switchable clutch (58, 59) for coupling to a drive shaft of the associated fuel pump (50, 52).

3. Drive stage according to claim 1 or 2, characterized in that the clutches (33"'A, 33"'B, 58, 59) assigned to each motor (33'; 33'A, 33'B) are switchable alternately.

4. Drive stage according to claim 2 or 3, characterized in that the respective motor (33'; 33'A, 33'B) has an output shaft (33"A, 33"B) which is connected or connectable at its first end to the first switchable clutch (33"'A, 33"'B) and at its second end to the second switchable clutch (58, 59).

5. Drive stage according to any one of the preceding claims, characterized in that the respective motor (33'; 33'A, 33'B) is an electric motor and that at least one energy storage device (62) is provided for storing electrical energy and supplying the respective motor (33'; 33'A, 33'B) with electrical energy. 6.Drive stage according to claim 5, characterized in that at least some of the motors (33'; 33'A, 33'B) can be operated in a generator mode in which electrical energy can be generated when coupled with an associated rotor drive and driven by an associated rotor arrangement (33, 33A, 33B), and that the motors in question (33'; 33'A, 33'B) are designed to feed the generated electrical energy back into the respective associated energy storage device (62). Drive stage according to any one of the preceding claims, characterized in that the power output of the respective motor (33'; 33'A, 33'B), in particular via a speed control, is controllable or adjustable both when driving the associated rotor arrangement (33, 33A, 33B) and when driving the associated fuel pump (50, 52).

8. Launch vehicle with at least one propulsion stage according to one of the preceding claims.

9. Method for operating a launch vehicle (1) according to claim 8, wherein the propulsion stage (3) is driven by the at least one reaction propulsion unit (36, 36A, 36B) during an ascent phase, wherein the propulsion stage (3) descends unpowered to Earth in a descent phase after the end of the ascent phase, and wherein the propulsion stage (3) is transferred into a landing phase shortly before reaching the Earth's surface, in which it steers towards a landing position in controlled flight and lands in a controlled manner, characterized in that at least some of the engines (33'; 33'A, 33'B) are coupled to an associated fuel pump (50, 52) during the ascent phase, and that at least some of the engines (33'; 33'A, 33'B) are coupled to an associated rotor assembly (33, 33A, 33B) during the landing phase, and the associated rotor assembly (33, 33A, 33B) controlled drive. 10.Method according to claim 9 for operating a launch vehicle (1) according to claim 8, characterized in that at least some of the motors (33'; 33'A, 33'B) designed as electric motors are coupled to an associated rotor arrangement (33, 33A, 33B) during the re-entry phase and are operated in a generator mode, wherein the associated rotor arrangement (33, 33A, 33B) drives the associated motor (33'; 33'A, 33'B) acting as a generator by autorotation of its rotors (34, 34'), and wherein the electrical energy thereby generated is fed back into an associated energy storage device (62) and stored there.

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