Archer Aviation Patente

Landing gear (10) for an aircraft (100), comprising a wheel (20) or a skid, a shock absorber (30) with an absorber connection (32) to connect to a fuselage (110) of the aircraft (100) and at least one attachment lever (40) with a lever connection (42) to connect to the fuselage (110) of the aircraft (100) spaced apart from the absorber connection (32), the wheel (20) or skid being attached to the attachment lever (40) spaced apart from the lever connection (42), wherein the shock absorber (30) is connected to the wheel (20) or skid directly or indirectly via a wheel or skid connection (34) to provide a load path (LP) for vertical load from the wheel (20) or skid to the absorber connection (32), and wherein the shock absorber (30) comprises a fuse pin (36) at the absorber connection (32) and/or the wheel or skid connection (34) having a predetermined breaking load to break and release the connection at the absorber connection (32) and/or the wheel or skid connection (34), and wherein the attachment lever (40) comprises a rotary bearing (41) providing a folding movement for the wheel (20) or skid from an extended landing position (ELP) into a crash position (CP) when the fuse pin (36) breaks.

Embodiments of the present disclosure provide systems and methods for averting, shedding, or otherwise managing ice accretions that may develop during flight of an aircraft. Example systems and methods selectively modulate propeller parameters in a way that does not disrupt a flight trajectory; direct oil from a lubrication and cooling path to targeted sections of ice-prone surfaces to manage ice accretion in a way that does not unduly increase the total volume of oil, require larger pumps, or complicate the system; or generate heat at targeted areas of a propeller assembly by electric heating systems that utilize propeller motion.

The present disclosure relates generally to control of systems with modular devices. In one embodiment, a device is disclosed, comprising: a plurality of electrical contacts; and at least one processor, wherein the plurality of electrical contacts are configured to be coupled to one of: a first voltage source or a second voltage source, wherein the at least one processor is configured to determine an identification (ID) value for the device by determining voltage values based on a coupling configuration of the plurality of electrical contacts.

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.

The invention is related to a ducted engine (10) for a vertical take-off and landing aircraft (100) having an engine inlet section (20) and an engine outlet section (40), and wherein the engine outlet section (40) comprises a moveable nozzle (42) being moveable between at least one cruise position (CP) and at least one hover position (HP), the engine inlet section (20) having an inlet opening area (IA) and the engine outlet section (40) of the nozzle (42) having a cruise opening area (CA) with the nozzle (42) in cruise position (CP) and a hover opening area (HA) with the nozzle (42) in hover position (HP), wherein the hover opening area (HA) is larger than the cruise opening area (CA), and wherein the hover opening area (HA) has a hover relation (HR) to the inlet opening area (IA) between 0.8 and 1.5 and the cruise opening area (CA) has a cruise relation (CR) to the inlet opening area (IA) between 0.5 and 1.2.

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.

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 relates to a multi-layer assembly, comprising at least one erosion barrier layer, at least one thermal insulation, at least one composite structure layer, and at least one gas barrier layer.

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 present invention relates to an acoustic liner (100) of a propulsion unit (500) of a vertical take-off and landing aircraft (600), which is operable between a hover position (HP) and a cruise position (CP). The acoustic liner (100) comprises a face-layer (110) with openings (111), a core-structure (120), which is coupled to an inner surface (113) of the face-layer (110) and comprises cells (121) with cell walls (122) surrounding cell cavities that are in fluid communication with the openings (111) to function as acoustic resonators. The acoustic liner (100) further comprises a base-layer (130) that covers the core-structure (120) on a side opposite to the face-layer (110). The core-structure (120) comprises also a plurality of drainage slots (140, 141, 142) that cut through the cell walls (122) of adjoining cells (121) for forming a network of interconnected fluid communication paths between the respectively slotted cells (121). The plurality of drainage slots (140, 141, 142) comprises a group of hover slots (141) and a group of cruise slots (142), which extend along slot directions (1410, 1420) that are different to each other. The acoustic liner (100) also comprises at least one drainage opening (131) that is in fluid communication with at least one of the slotted cells (121) for draining fluid entering the cell cavities via the openings (111) through the drainage slots (140, 141, 142). The present invention also relates to a propulsion unit (500) and an aircraft (600)

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).

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).

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).

The present invention relates to a bladed disk (100) for a driven rotor of an engine of an aircraft propulsion unit, the bladed disk comprising: a hub (102), a plurality of blades (104) extending radially from the hub, and at least one retaining element (200a-c), wherein the at least one retaining element extends circumferentially to the hub (102) and is coupled to the hub to retain a segment of the bladed disk (100) in the event of a failure of the bladed disk. The present invention also relates to a propulsion unit for an aircraft, the propulsion unit comprising an engine comprising the bladed disk (100), and an aircraft comprising such a propulsion unit.

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 propulsion unit (10) for an aircraft, the propulsion unit comprising: a pair of engines (100a, 100b), each engine having a bladed disk driven by a motor; and a containment ring (20) located downstream of an air intake of the engines, wherein the containment ring (20) comprises a main body having at least two openings (203a, 203b) to provide air flow paths to the pair of engines, and wherein the main body comprises an outer part (200) circumferential to the at least two openings to protect components adjacent to the propulsion unit against damage in a case of blade failure, the main body further comprising a separating part (202) with a reduced mechanical stability compared to the outer part to separate the openings from each other. The present invention also relates to an aircraft, preferably a vertical take-off and landing aircraft, comprising the propulsion unit (10).

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 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 stator body (10) for a rotating electric machine, comprising: a ring-shaped outer stator yoke (11); a plurality of stator teeth (13) and a plurality of coil pockets (14). According to the invention, a plurality of receptive cavities (12) is recessed into the stator yoke (11) for receiving a plurality of heat sink means (20) for heat dissipation on an outside of the stator body (10); wherein the heat sink means (20) comprise at least one heat flux material having a higher thermal conductivity than a magnetic flux material of at least the stator yoke (11) of the stator body (10).

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 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 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.

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 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).

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).

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