Lilium Patente

The invention relates to a wing spar structure 210 of an aircraft 100. The wing spar structure comprises a left spar 214 associated to a left wing 202b of the aircraft and a right spar 212 associated to a right wing 202a of the aircraft. Each spar 212, 214 has a web portion 230 extending in a substantially vertical direction v and at least one spar cap 232, 232a, 232b extending in a substantially orthogonal to the web portion 230. According to a first aspect of the invention, said left spar 214 is connected to said right spar 212 by connecting the spar cap 232b of the left spar 214 to the spar cap 232a of the right spar 212 at a spar connection portion 240, 240a, 240b. According to a second aspect of the invention, said left spar 214 is connected to said right spar 212 in a fuselage region 224 at a spar connection portion 240, 240a, 240b and a height 234 of the web portion 230 in vertical direction v increases towards the spar connection portion 240, 240a, 240b.

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The present invention relates to an aircraft (10), comprising a fuselage (12), at least one pair of wings (14, 16) and a battery system, wherein the battery system comprises at least one battery pack (24), each battery pack (24) comprising a number of individual battery modules (26), which are directly or indirectly coupled to one another. The at least one battery pack (24) can be disposed between an inner structural wall (20) defining an interior space (18) of the fuselage (12) and an outer fairing wall (22) of the fuselage (12). The fuselage (12) can be provided with a rack mounting mechanism (40) comprising a number of mounting brackets (42), each for exchangeably mounting one of the battery modules (26) to the aircraft (10). Each of the battery packs (24) can be a virtual battery pack (24), which is obtained by electrically connecting a predetermined number of the battery modules (26).

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The present invention relates to an aircraft (10), comprising a fuselage (12), at least one pair of wings (14) and a battery assembly for providing power to electrical systems of the aircraft (10), wherein the battery assembly comprises a number of individual battery modules (18) which are directly or indirectly coupled to one another, the fuselage (12) is provided with a mounting assembly (16) with a number of mounting positions (16a, 16b, 16c) for each holding one of the battery modules (18), and the number of mounting positions (16a, 16b, 16c) is larger than the number of battery modules (18) such that in a mounted state of all battery modules (18), at least one of the mounting positions (16a, 16b, 16c) remains vacant (20) thus defining a placement configuration of the battery modules (18) and the vacant mounting positions (20), and/or the mounting assembly is provided with at least one displacement assembly which allows to displace at least one of the battery modules with respect to the fuselage.

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Provided is a solid state power controller (SSPC). The SSPC comprises a power supply line configured to be connected between a power source and a load. Further, the solid state power controller comprises a semiconductor switching unit provided on the power supply line and configured to switch between at least two states according to a command signal. The at least two states include a conducting state and a non-conducting state. Further the solid state power controller comprises a state machine, which is configured to exhibit at least two states, including an ON state and an OFF state. The state machine is further configured to output the command signal according to a current state of the state machine.

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Valve apparatus for controlling fluid flow through a duct of an air conditioning system of an aircraft, the valve apparatus comprising:• a valve seat body (2) having a valve seat hole (3) being configured to allow the fluid to flow through it,• a damper body (4) for closing the valve seat hole (3),• wherein the valve seat body (2) and the damper body (4) are movable relative to each other between an open configuration and a closed configuration, wherein in the open configuration the fluid may flow through the valve seat hole (3) and in the closed configuration the damper body (4) interacts with the valve seat body (2) in a sealing manner such that the fluid may not flow through the valve seat hole (3),• wherein the valve seat hole (3) has a hole contour formed by a hole edge (5) and the damper body has a damper contour formed by a damper edge (6), the hole edge and the damper edge being configured to interact with each other in the sealing manner thereby forming a circumferential sealing line having a circumferential sealing length,• wherein the hole contour has a plurality of hole protuberances (7) and the damper contour has a plurality of damper protrusions (8), so that the circumferential sealing length is larger than a reference length being defined by a length of a circumference of an equal area circle (EAC) having an area equal to an area of the valve seat hole (3),• wherein the valve apparatus comprises a solenoid actuator (9) having an actuator rod for transmitting an actuator force, the solenoid actuator being mounted for moving the damper body (4) relative to the valve seat body (2) towards the closed or open configuration, and• wherein force transmission means (11) are arranged between the actuator rod (10) and the damper body (4), the force transmission (11) means being configured to apply a closing force to the damper body that is larger than the actuator force.

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An engine (100) for an aircraft is presented comprising a rotor (101), an electric motor (103) actuating the rotation of the rotor (101) and an ECU (104) controlling the electric motor (103), wherein the rotation of the rotor (101) provides a main flow of air causing the thrust of the engine (100) and wherein the ECU (104) is located within the volume defined by the main flow.

The present disclosure relates to an actuator (1) for generating an actuating movement for a primary flight control surface (2). This actuator (1) comprises an electric motor (5) and a reduction gear coupled thereto. This reduction gear is formed as a harmonic drive gear (strain wave gear) (6). The disclosure also relates to a civil aircraft equipped with such an actuator (1). In this aircraft, the actuator (1) according to the disclosure can be used in particular as a flap actuator for adjusting a flap structure (2) pivotally hinged to a trailing end of a wing structure (3).

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The present invention provides an engine 310 of a vertical take-off and landing aircraft 300, wherein the engine is configured to be movable with respect to an aircraft component 342 of the aircraft 300 between a hover position for take-off and landing, and a cruise position for forward flight, wherein the engine 310 comprises an aerodynamic component 332 having at least one aerodynamic element 334 movable between a first position 336 according to a first operational state of the aircraft, and a second position 338 according to a second operational state of the aircraft, the aerodynamic element defining an aerodynamic surface in contact with an airstream passing through the engine.

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The present invention relates to a battery management system and method for performing a battery state and optionally health parameter observation, in particular cell state-of-charge (SOC) observation, with two redundant, independent and dissimilar lanes. Specifically, a SOC determination in a first one of the lanes is based on Coulomb Counting. The other lane employs a different algorithm than Coulomb Counting. In embodiments, a battery health observation is further performed independently by the two lanes, wherein the first lane employs an ageing model and the other lane a different (dissimilar) algorithm. On the basis of state and health observation, state (state of function) of the battery system can be predicted to determine a range of flight in accordance with a predetermined flight profile.

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