Lilium

The present invention relates to a casing wall structure (100) for a battery assembly (10), wherein the battery assembly (10) comprises at least one stack (12) of battery cells (14), and a casing with a base plate (16) and a top plate (18), the casing wall structure (100) comprising an inner wall (102) facing the battery cell stack (12), an outer wall (104) facing to the outside of the battery assembly (10), and a cold plate (106) which is sandwiched between the inner wall (102) and the outer wall (104), wherein the inner wall (102) is at least partially formed with a smaller thickness than the outer wall (104). The invention further relates to a battery assembly comprising such at least one such casing wall structure and an electrical propulsion aircraft comprising at least one such battery assembly.

The present invention relates to an electrical propulsion aircraft (100), comprising a fuselage (102) for hosting payload with a tail section (104), a center section (106) and a nose section (108), a pair of wings (100) attached to the fuselage (102) for providing lift to the aircraft (100), an empennage (112) attached on the tail section (104) of the fuselage (102) for providing stability during flight of the aircraft (100), an electric propulsion system for providing thrust to the aircraft, comprising a plurality of battery modules (114) for storing electrical energy, a plurality of propulsion units (118, 118a) attached to the empennage (112) or the tail section (104) of the fuselage (102), wherein each propulsion unit (118, 118a) comprises an electric motor for driving a propulsor. According to the invention, the empennage (112) comprises a horizontal (112a) stabilizer provided with at least one of the propulsion units (118) on each of its sides with respect to a longitudinal centerline (C) of the aircraft (100) for providing thrust and directional control of the aircraft (100).

The invention relates to a blade (10) for a fan (20) of a ducted engine (30) of an aircraft (100), the fan (20) being located in an engine duct (34) between an engine inlet section (32) and an engine outlet section (36), wherein the blade (10) comprises a blade body (12) having an aerodynamic shape to generate a propelling force and extending from a blade root (14) to a blade tip (16), and wherein the blade body (12) comprises an aerodynamic profile (AP) comprising a leading edge (LE) and a trailing edge (TE) defining a chord line (CHL) and further comprising an upper profile (UP) and a lower profile (LP) defining a camber line (CAL), and wherein the tangent (T) at the leading edge (LE) to the camber line (CAL) defines a leading edge direction (LED) for the leading edge (LE), the leading edge (LE) having an incidence angle (IA) to a flow direction (FD) of the ducted engine (30), characterised in that the blade body (12) comprises a tip section (17) at the blade tip (16) having the aerodynamic profile (AP) with the tip incidence angle (TIA) being smaller than an optimum incidence angle (OIA) of the aerodynamic profile (AP).

An inertial sensor unit (1, 1a, 1b, 1c) comprising: an inertial sensor (2, 2a, 2b, 2c), for detecting an inertial quantity (a₀, Ω₀), in particular acceleration and/or an angular rate, the inertial sensor having a sensitive unit (21, 21b) for sensing the inertial quantity (a₀, Ω₀) and a signal processing unit (22, 22b, 22c) being configured to process an output of the sensitive element and to output a signal (a, Ω) related to the sensed inertial quantity; and a virtual inertial quantity input unit (3, 3b, 3c) which is configured to input a virtual inertial quantity (a*, Ω*) to the inertial sensor (2, 2a, 2b, 2c, 2d). In order to improve hardware-in-the-loop suitability, the signal processing unit (2, 2a, 2b, 2c, 2d) is configured to output the signal based on the virtual inertial quantity (a*, Ω*) input by the virtual inertial quantity input unit (3, 3b, 3c).

The invention is related to a Method for generating control limit outputs (CLO) for controlling an aircraft (100), comprising the following steps: - Determination of the present flight status (PFS) of the aircraft (100), - Calculation of at least one present control limit (PCL) for a controllable flight parameter (CFP) of the aircraft (100) corelated with the determined present flight status (PFS), - Providing a control limit output (CLO) based on the calculated at least one present control limit (PCL).

A battery assembly (10), in particular for use in an electrical propulsion VTOL aircraft (100), comprises a plurality of battery stacks (12), and a venting system (59) with a venting plenum (62), each battery stack (12) comprising a plurality of battery cells (50) stacked in a stacking direction (V) and a pressure relief safety device (30) configured to discharge gas from the battery stack (12) into the venting plenum (62) when a pressure inside the battery stack (12) exceeds a given threshold, and wherein the battery stacks (12) of the assembly (10) are arranged side by side along a row direction (R) perpendicular to the stacking direction (V), the pressure relief safety device (30) of each stack (12) being provided on a battery stack wall (19) extending in parallel to the stacking direction (V) and to the row direction (R), wherein the battery stack walls (19) are not formed as separate components but as a common continuous wall (21) attached to and connecting all the battery stacks (12) of the assembly (10), and wherein the venting system (59) is a venting cover (60) sealingly attached to and covering the common continuous wall (21).

The invention relates to an aircraft (1) with an electric propulsion system (2), wherein the electric propulsion system (2) comprises engines (21) and battery assemblies (22) to provide the engines with electric energy, a number of skin panels (45) are attached to frames (43) and stringers (42) of a fuselage (4) and defining the outer shape of the fuselage (4) and of which skin panels (45) covering holding bays (44) for the battery assemblies (22) to the outside of the fuselage (4). To improve the security of the aircraft in regard to a runaway event of the electric propulsion system the skin panels (45) are providing venting ports (60) which are connected to one or more venting manifold (11). The invention further relates to a venting manifold assembly (10) for an aircraft (1) with an electric propulsion system (2) having both an inlet mounting adapter (13) and an outlet mounting adapter (15) to be connectable to a housing (24) of a battery assembly (22) inside the fuselage (4) and a skin panel (45) of the aircraft (1), thus, increasing the durability and/or applicability of the venting manifold assembly as well as reducing the production cost of the venting manifold assembly.

The invention is related to a charging method for charging multiple battery packs (12) of an electrical power system (10), wherein each battery pack (12) comprises a switch (22) for switching between a charging state and a noncharging state of the battery pack (12), wherein the charging method comprises the following steps:measuring a voltage of each battery pack (12) and determining a lowest voltage battery pack (12),connecting the lowest voltage battery pack (12) to a charger with the switch (22) of the lowest voltage battery pack (12) and requesting a charging current from the charger,measuring the voltage of the lowest voltage battery pack (12) while charging the lowest voltage battery pack (12), andconnecting a next higher voltage battery pack (12) to the charger in parallel to the lowest voltage battery pack (12) with the switch (22) of the next higher voltage battery pack (12) if the voltage of the lowest voltage battery pack (12) being charged is within a predefined safe range of a voltage of the next higher voltage battery pack (12) and requesting an amended charging current from the charger for charging the lowest voltage battery pack (12) and the next higher voltage battery pack (12) together.

The present invention relates to a battery pack stack (10), comprising multiple pouch-type battery cells (12), a pouch, at least one positive and one negative electrode tab (12a, 12b) provided to corresponding electrode plates, and a cooling device (20) which is adapted to cool the battery cells (12), comprising a plurality of heat transfer pads (36), each of which is applied to one of the electrode tabs (30, 32) inside and/or outside the pouch, and a linking structure (38) interconnecting the heat transfer pads (36) along the stacking direction (S) of the electrode assembly.

The present invention relates to a battery assembly (10), comprising a battery cell stack (12) with a plurality of battery cells (12a) with respective electrode tabs (14), wherein the battery cells (12a) are stacked and electrically interconnected via their electrode tabs (14) along a stacking direction (S), and a housing with a top and a base plate, between which the battery cell stack (12) is compressed, and structural enclosure walls (16), wherein at least one barrier element (18) is included in the battery cell stack (12) between a pair of neighbouring battery cells (12a) for dividing the battery cell stack (12) into a plurality of sealed cell modules (20a. 20b), wherein the at least one barrier element (18) comprises at least one stud element (22), which extends through an opening (16a) in a first of the structural enclosure walls (16) and is fixed to the first enclosure wall (16) by means of a counter element (22a), wherein between the at least one barrier element (18) and the first enclosure wall (16), a busbar assembly (24) is provided for electrically inter-connecting the neighbouring cell modules (12a), wherein by means of the barrier element (18), the first enclosure wall (16), the busbar assembly (24) and an insulating sealing layer (26) arranged between the busbar assembly (24) and the first enclosure wall (26), a hermetic sealing of the neighbouring cell modules (20a, 20b) is provided.