Lilium Patente

The invention is related to a Method for producing a battery device (10), in particular for use in an electric aircraft, comprising the following steps:- Providing a battery housing (20) with at least one cell compartment (30) having one receiving opening (32) and surrounding compartment walls (34) for holding a battery module (40) with multiple battery cells (42),- Stacking multiple battery cells (42) on top of each other to create at least one battery module (40),- Inserting the at least one battery module (40) into the at least one cell compartment (30) through the receiving opening (32),- Closing the receiving opening (32) of the at least one cell compartment (30).

The invention is related to a Battery housing (10) for accommodating a battery device (20) with multiple battery cells (22) for an electric aircraft (100), comprising at least one inner wall (30) and at least one outer wall (40) forming together a battery volume (50) for holding the multiple battery cells (22), characterised in that the at least one outer wall (40) comprises an outside surface (42) wherein the outside surface (42) has an aerodynamical shape for completing a part of a fuselage (110) of the electric aircraft (100).

The present invention relates to a method (10) of detecting engine stall in an engine (1100) of a vertical take-off and landing aircraft (2000) under flight conditions (FC). The engine (1100) comprises a rotor (1130). In the method (10), a rotor torque (M_rotor) and a rotor speed (n_rotor) of the rotor (1130) are acquired. Based thereon, a stall ratio (SR) between the acquired rotor torque (M_rotor) and rotor speed (n_rotor) is determined. The determined stall ratio (SR) is compared with a stall boundary (SB) for the engine (1100), which is based on the acquired rotor torque (M_rotor) and rotor speed (n_rotor), for detecting occurring engine stall. The result of the comparison is outputted as a detection result (DR). The invention relates further to a computer program product, a control system (100), a propulsion system (1000) and an aircraft (2000) utilizing and implementing said method (10).

Welding preparation device (1) for preparing at least two battery cell tabs (P1, P2) for welding comprising: a clamping mechanism (2) comprising: a support part (3); and a clamping part (4) for clamping the at least two battery cell tabs (P1, P2) at a clamping force exertion portion (41) thereof together with the support part (3) therebetween within a clamping region (CR) by a clamping force along a clamping force direction. In order to improve welding quality, the clamping mechanism (2) has an access portion (7) which, in a clamped state, allows access to the clamping region (CR) in order to provide a weld seam to the battery cell tabs (P1, P2).

This patent application introduces a wing assembly for aircraft, featuring an innovative approach to integrating a servo actuator control unit (SCU) (5) within the propulsion unit (2). The assembly includes a main section for attachment to the aircraft's fuselage and a pivotally connected flap section (1). The propulsion unit (2), integral to the flap section (1), is configured with inlet and exhaust areas (2a, 2b) to facilitate thrust production. The SCU (5) is uniquely integrated into a skin part (2c) of the propulsion unit (2), positioning it in the path of airflow generated during flight. This strategic placement enhances heat dissipation by utilizing the airflow, effectively reducing the need for additional cooling mechanisms and contributing to a lighter, less complex aircraft design. The direct exposure of the SCU (5) to airflow helps maintain it within safe operational temperature ranges, potentially increasing its reliability and service life. The configuration focuses on improving thermal management and aerodynamic efficiency, aligning with current trends in aerospace for more efficient, robust, and environmentally conscious aircraft wing assemblies.

The present invention provides an aircraft with an electric propulsion system, a fuselage (10) configured to host payload and energy storage elements (20), and lift generating wings attached to the fuselage (10). The fuselage (10) comprises an aerostructure to bear and distribute loads between the fuselage (10) and the wings. The aerostructure comprises a structure of frames (30), stringers and skin panels (41, 42, 43). A plurality of frames (30) are separated in a longitudinal direction of the fuselage (10) and arranged in parallel to each other, defining holding spaces (22) in between adjacent frames (30). A plurality of stringers extending predominately in the longitudinal direction of the fuselage (10) are separated in a circumferential direction of the fuselage (10) and are attached to a number of the frames (30). A plurality of skin panels are attached to the frames (30) and stringers, wherein a first (41) and second (42) set of the skin panels are defining the outer shape of the fuselage (10), the second set of the skin panels (42) covering the holding spaces (22) to the outside of the fuselage, and a third set of the skin panels (43) is covering and separating an interior of the fuselage (10) from the holding spaces (22). The holding spaces (22) are configured to host the energy storage elements (20), which are attached at least to the frames (30) and/or stringers of the aerostructure.

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The invention refers to an aircraft (2), particularly an aircraft (2) capable of vertical take-off and landing, comprising a load carrying structure (4), and a variable lift body (6) defining an aerofoil (55) and being moveably attached to the load carrying structure (4), wherein the variable lift body (6) is pivotable around a first axis;wherein a rotary actor is adapted to cause the variable lift body (6) to pivot in relation to the load carrying structure (4) arranged within the aerofoil (55).

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The invention provides a liner for a cargo and/or accessory compartment (2) of an aircraft, wherein the liner is fire-resistant and gas tight to the extent of fire starving capability and keeps smoke emissions to a neglectable level in terms of flight safety, the liner comprising a bag (3) of fire-resistant and gas tight flexible fabric. The invention also provides an aircraft with such a liner. Furthermore, the invention provides a method of lining a cargo and/or accessory compartment of an aircraft with a liner, comprising the steps of preassembling a liner according to the present invention, folding the preassembled liner, placing the folded liner inside a fuselage structure of the aircraft, unfolding the liner and attaching the liner to the fuselage structure of the aircraft.

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An inceptor apparatus (30a; 30b) for an aircraft comprises a primary inceptor member (32a; 32b) provided in the form of a stick member having a grip portion, at which the stick member can be gripped by a pilot's hand, and a secondary inceptor member (50a; 50b) provided at an upper portion of the primary inceptor member (32a; 32b) and having an actuating portion, at which the secondary inceptor member can be manually actuated by a pilot's thumb. Both inceptor members (32a, 50a; 32b, 50b) have associated a respective sensor assembly which is provided to generate electronic flight control signals or commands in response to at least one of i) pivoting movements of the respective inceptor member (32a; 32b; 50a; 50b) around each of two independent maneuvering axes associated to the inceptor member, ii) forces acting on or via the respective inceptor member in pivoting directions with respect to each of the maneuvering axes, and iii) lateral flexing or bending of the respective inceptor member. According to one aspect of the invention, a flight control computer system (12) for an aircraft has at least one such inceptor apparatus (30a, 30b) and is configured to effect the flight control in a primary control mode based on flight control signals or commands corresponding to or being based on the electronic flight control signals or commands generated by the primary sensor assembly (40a; 40b) of the inceptor apparatus (30a; 30b) and to effect the flight control in a secondary control mode based on flight control signals or commands corresponding to or being based on the electronic flight control signals or commands generated by the secondary sensor assembly (62a; 62b) of the inceptor apparatus (30a; 30b).

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An electrical power distribution network (306) of an electric power system (300) of an aircraft is operated such that it sequentially adopts a plurality of different partial load sharing modes in a time variable manner, which provide for partial load sharing across electrical power sources (A, B, C, D) with respect to associated electrical loads (AA, BB, CC, DD), by sequentially switching between a plurality of different partial load sharing configurations of the electrical power distribution network, each partial load sharing configuration being associated to a particular one of the partial load sharing modes.

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The present invention relates to a battery management system and method for performing a battery health parameter observation, in particular cell impedance observation, with two redundant, independent and dissimilar lanes. Specifically, a cell impedance observation in a first one of the lanes is based on Electrochemical Impedance Spectroscopy, EIS. The other lane employs a different algorithm than EIS. In embodiments, a battery state observation is further performed independently by the two lanes, wherein again the first lane employs EIS 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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An electrical power distribution network (306) of an electric power system (300) of an aircraft is operated in at least one normal operation mode such that it provides for load sharing across electrical power sources (A, B, C, D) with respect to electrical loads (AA, BB, CC, DD), wherein the electrical power distribution network (306), in case of an electrical fault, is operated in at least one electrical failure mitigating operation mode, which provides for electric fault isolation, such that a network portion of the electrical power distribution network (306) including the electrical fault is isolated from at least one other network portion of the of the electrical power distribution network.

A wing (1) comprises: a fixed lift generating portion (2) extending along an extension direction of the wing (1); a moveable lift modifying portion (3) which is moveably, preferably pivotably, more preferably about a pivot axis (P) along the extension direction of the wing (1), connected to the fixed lift modifying portion (2), wherein the moveable lift modifying portion (3) comprises at least one propulsion device (31) configured to generate a thrust force along a thrust direction for powering the aerial vehicle, preferably a plurality of propulsion devices (31) arranged along the extension direction. In order to enhance the aerodynamic performance, the at least one propulsion device (31) forms a moveable wing tip end portion (4) located at a tip side along the extension direction of the wing (1), the moveable wing tip end portion (4) comprising an aerodynamic profile, which is in particular part of a winglet.

A flight control system (10) for an aircraft comprises a flight control computer system (12), which is connected via an electronic or optoelectronic bus system (22) with a plurality of bus nodes (14, 16, 18, 20), which each are configured to at least one of controlling an associated aircraft device based on command messages received from the flight control computer system via the bus system and sending information messages to the flight control computer system via the bus system. According to one aspect of the disclosure, the electronic or optoelectronic bus system is a redundant electronic or optoelectronic bus system (22) comprising plural independent bus sub-systems (22a, 22b), wherein each bus node is configured to communicate with the flight control computer system (12) via two different bus sub-systems (22a, 22b) of the plural independent bus sub-systems, wherein each bus node further is configured to communicate with the flight control computer system on basis of an associated predetermined bus communication protocol via a first bus sub-system (22a) of the respective two different bus sub-systems and on basis of an associated predetermined bus communication protocol via a second bus sub-system (22b) of the respective two different bus sub-systems.

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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).

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.