Archer Aviation

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 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 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 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 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).

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)

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.