49results about "Emergency apparatus" patented technology

The composite sheet includes a textile component having a first surface and a second surface, and a first polymer coating on the first surface of the textile component, the textile component further including: a first structure of at least one woven fabric of continuous filament yarns, the woven fabric having an areal weight of about 70 to about 508 gsm; a second structure of a nonwoven fabric comprising a blend of discontinuous fibers of oxidized polyacrylonitrile fibers and silica fibers, the oxidized polyacrylonitrile fibers being present in an amount of about 30 to about 70 weight percent of the combined weight of the oxidized polyacrylonitrile fibers and the silica fibers; and a third structure of at least one woven fabric of continuous filament yarns, the woven fabric having an areal weight of about 70 to about 508 gsm, wherein the nonwoven fabric of the second structure may be mechanically attached to at least one of the other textile structures by an attachment means.

The present application provides a fall damage reduction system, a flying device, and a vehicle-mounted device. The flying device includes a flight communication section (24), a device fall determination section (25, S101), and a flight notification section (25, S104, S105). The flight communication section performs wireless communication in accordance with a communication standard for inter-vehicle communication. The device fall determination section determines whether the flying device is likely to fall. In a case where the flying device is determined by the device fall determination section to be likely to fall, the flight notification section transmits fall information including information indicating a presumed fall position via the flight communication section. The vehicle-mounted device includes a vehicle communication section (34) and a processing section (37, S207, S208, S210, S211, S212, S214, S215, S218, S219). The vehicle communication section performs wireless communication in accordance with a communication standard for inter-vehicle communication. In a case where the fall information is received via the vehicle communication section, the processing section executes fall avoidance processing that is processing for avoiding the fall position.

The invention relates to a method and system for deflating an oxygen mask. The method for throwing the oxygen mask comprises the steps that the current flight scene S of an aircraft is determined, wherein the flight scene S comprises a ground station-crossing and take-off stage S1, a climbing or cruising stage S2 and a descending and landing stage S3; calculating a cabin height warning value ZCW based on the flight scene S and the cabin height value ZC; judging a jettisoning height threshold value ZD based on the cabin height alarm value ZCW; in any flight scenario, the cabin height value ZC is compared to the jettisoning height threshold ZD, and when the cabin height value ZC is greater than the jettisoning height threshold ZD, the oxygen mask is automatically jettisoned.

An electrical device for protecting an electrical installation intended to be placed between a power supply and an electrical load, the device including: a power contactor comprising two contacts that are intended to be connected to the power supply; at least two Hall sensors configured to measure the electric currents at the input of the power contactor; an electrical precharging circuit intended to be placed between the power supply and the power contactor and connected to the electrical load; a relay configured to connect or disconnect the electrical precharging circuit; a pyrotechnic electrical cutoff device intended to be connected to the power supply and to the power contactor, and a controller configured to receive the measurements from the Hall sensors and to open or close the power contactor and the relay.

An embodiment aircraft includes a fuselage, a wing coupled to the fuselage, a fixing portion connecting a portion of the wing to the fuselage, and a controller disposed in the fuselage, wherein the controller is configured to transmit a severing signal to cause a portion of the fixing portion to be severed.

A light unit executes up to three different functions (cabin area illumination, individual / dedicated / decorative illumination and emergency illumination) and an emergency illuminated sign unit, each one with an internal controller and a rechargeable capacitor. An example non-limiting embodiment also provides a cabin light system and an emergency lighting system, where each illumination unit (light source or illuminated sign) is as described above.

A method of guidance for an aircraft includes obtaining relative position data indicative of a direction from a location of the aircraft to a location of a moving target point. The method also includes determining, based on the relative position data, a location of a virtual leader based on projecting the location of the aircraft onto a virtual leader path, where the virtual leader path corresponds to a straight approach path to the location of the moving target point along a desired approach direction. The method further includes determining, based on the location of the virtual leader, a guidance input to cause the aircraft to follow the virtual leader.

An electromechanical actuator assembly (EMA) for use with a reclinable passenger seat. The EMA includes a housing, an electric motor, and an epicyclic gearbox including first and second ring gears. In the event of power loss, the second ring gear can be decoupled from the first ring gear to permit relative rotation, for instance to manually return the reclined seat to upright, by a mechanical override mechanism including first and second spring-loaded shoot bolts coupled to the first ring gear, a displaceable plate engaging one end of the first and second spring-loaded shoot bolts, and a manual release pin coupled to the displaceable plate. In use, the manual release pin is moveable axially to a position in which the displaceable plate displaces toward the first ring gear to decouple the first and second ring gears.

Aircraft portion comprising a depressurization-safe equipment door. The aircraft portion comprises equipment including: - an outer wall delimiting an enclosure; - a door facing an aisle and movable between a closed and an open configuration; wherein the aisle has: - a maximum width when the door is in the closed configuration; - a minimum width when the door is in the open configuration; and wherein, when the door is in the closed configuration, and in the event of depressurization, the door is moved and maintained in an emergency configuration in which the aisle has an optimal width that is sufficient to simultaneously: - allow air to escape from the enclosure; and - allow passengers to use the aisle and pass in front of the equipment. Figure for abbreviation: Figure 4

The present invention relates to a method for controlling an aircraft (1) comprising at least two engines (16, 17), each engine (16, 17) being arranged in an engine compartment (21, 22) of its own, each engine (16, 17) being connected to a fuel supply circuit (31, 36) provided with a supply cut-off switch (34, 38) of its own for cutting off a fuel supply, said aircraft (1) comprising at least one fire detector (26-29) in each engine compartment (21, 22), said aircraft (1) comprising at least a first fire extinguisher (40) and a second fire extinguisher (45).An automatic assistance phase (PHASASSIST) comprises, in the event of a fire detected in an engine compartment, automatically and with an avionics system (60), an activation of the power cut-off switch (34, 38) of the engine (16, 17) present in the compartment concerned, then a triggering of the first extinguisher (40), then a triggering of the second extinguisher (45) under respective predetermined conditions.

This vehicle comprises: a front seat 2 the posture of which can be changed between a first posture in which a seating surface faces forward and a second posture in which the seating surface faces rearward; a rear seat 3 that is arranged behind the front seat 2 and that has a seating surface facing forward; and a floor airbag module 4 that is provided between the front seat 2 in the second posture and the rear seat 3, and that has a floor airbag which emerges between the front seat 2 and the rear seat 3 if an impact force acts on a host vehicle or an impact force is expected to act on the host vehicle, .

Aspects of the present disclosure relate to flight control of electric aircraft and other vehicles. In one embodiment, an aircraft is disclosed which comprises a fuselage, one or more flight control computers configured to provide control signals, four sets of electrically driven propellers, the first set being positioned in front of the second set and both on one side of the fuselage, the third set being positioned in front of the fourth set and both on the other side of the fuselage, and a plurality of electric buses, the one or more flight control computers configured to provide control signals to at least one of the first set of propellers, at least one of the fourth set of propellers, and at least one of the third set of propellers and at least one of the second set of propellers via one of the plurality of electric buses.

A barrier rope assembly which can be disassembled on one side, the barrier rope assembly being provided in a door frame of an aircraft having opposite first and second door frames, and including a barrier rope and a hook for hooking the barrier rope. The barrier rope assembly includes a support member which is located on the side of the first door frame opposite the second door frame and supports one end of the barrier rope, and a connecting member which connects the support member to the first door frame in such a manner that the support member can be disassembled from the side of the second door frame relative to the first door frame.

An embodiment aircraft includes a fuselage, a wing coupled to the fuselage, a fixing portion connecting a portion of the wing to the fuselage, and a controller disposed in the fuselage, wherein the controller is configured to transmit a severing signal to cause a portion of the fixing portion to be severed.

An emergency vision assurance system device includes a housing; an inflatable enclosure having an inflated state and a deflated state, the main enclosure in the deflated state being disposed in the housing; a blower disposed in the housing; a hose with a first end attached to the blower and second end attached to the enclosure; and a switch disposed at the second end for operating the blower. The emergency vision assurance system device also includes a housing with an opening for access to the inflatable enclosure, the opening having a smaller opening when closed and a larger opening when open; and a cover for closing the opening.

Device for ventilation, comprising: a breathing gas source for providing breathing gas for the ventilation of a patient; at least one sensor connected to an interface (8) of the device, which can be positioned on the patient's skin and is designed to non-invasively measure at least one component of the patient's blood as at least one ventilator-dependent blood parameter; a control device designed to evaluate at least one blood parameter and autonomously change at least one operating parameter of the device depending on the at least one blood parameter, wherein maximum corridors are specified for the at least one operating parameter to prevent implausible automatic ventilation settings, and operator inputs are evaluated to take external control specifications into account; a device housing (1) with a control panel (2) and a display (3), wherein an oxygen supply valve is adaptable to the device in the area of ​​the device housing (1) and a respiratory gas pump is arranged in an interior of the device, wherein a humidifier is also adaptable to the device, wherein the display (3) is designed as a display so that a physician can select a disease pattern directly on the display; a modem or other interface for transmitting recorded data to the physician, as well as abnormalities, operating hours, or other data to a maintenance or customer service department to ensure proper functioning of the device; wherein the control device is designed to change the at least one operating parameter within a control loop, to which data of the at least one blood parameter are supplied as setpoint and actual values, and to assign strategies for carrying out the control to specific disease patterns, wherein each strategy includes a priority list of 1 to N different settings that are to be processed within the respective bandwidths when applying the strategy, wherein the strategies with their respective settings and bandwidths can be stored and made available as fixed predefined strategies for specific disease patterns; wherein the control device is further configured to take into account characteristic settings and their ranges for the disease pattern chosen by the physician and to apply the settings taken into account; wherein at least one blood parameter includes a carbon dioxide partial pressure and the control device is further designed to compare an actual value of the carbon dioxide partial pressure with a setpoint when performing the control and, if a current strategy has been fully executed and the setpoint has not been reached, to automatically apply a new strategy suitable for further improving the patient's condition.

A device for protecting a hydrogen tank, in particular for an aircraft, comprising an external envelope delimiting an internal volume, and an extinguishing agent, preferably in the gas phase, in the internal volume, the external envelope being provided with at least one pressure relief valve, to release the at least one extinguishing agent when a pressure prevailing in the internal volume is greater than a threshold value, and the external envelope being shaped to be installed in a fuselage of the aircraft.

The invention belongs to the field of aviation, and relates to a cabin aircraft outer throwing cover starting mechanism. The device comprises a machine body opening cover, a cover throwing starting mechanism and a detonation firing mechanism. The cover throwing starting mechanism comprises an actuating pull rod, a limiting device, a steel cable winch, a fixed pulley, a steel cable, a cover throwing pull pin and a mounting support. In the process of throwing the cover outside the aircraft, the aircraft body opening cover is opened firstly, the steel cable which is bound and fixed to the actuating pull rod in advance is detached, then the steel cable is pulled, the actuating pull rod rotates to cut off the shear pin on the limiting device, then the throwing cover pull pin is driven to be pulled out, the detonation firing mechanism outputs a detonation detonation signal, and output of the throwing cover starting signal is completed.

Backup device (1) for a hybrid aircraft, comprising a turbomachine having at least one gas generator (12), a free turbine (14), and a main rotor (52), the device (1) comprising a first reversible electric machine (30) coupled to the main rotor (52) by means of a first and a second switchable coupling means (32, 42) mounted in parallel, operating in a motor mode so as to drive the main rotor (52) by means of the first switchable coupling means (32), and operating in a generator mode to generate electrical power by being driven by the main rotor (52) by means of the second switchable coupling means (42), the latter being deactivated when a rotational speed of the main rotor (52) is greater than or equal to a first predetermined stall threshold value (N1),such that the first electric machine (30) cannot be driven in generator mode by the main rotor (52). Fig. 2.

To provide a power supply circuit capable of supplying power to a load even when an excessive shock occurs to a mobile.SOLUTION: A power supply circuit includes a first circuit breaker 50a that switches between a conducting state and an interrupting state, and a second circuit breaker 56a that is provided in a second power transmission path 28 that sends power from a second power source 20 to a load 14, and switches between a conductive state and a cutoff state between the second power source 20 and the load 14. The first circuit breaker 50a includes a first fixed contact and a first movable contact, and the second circuit breaker 56a includes a second fixed contact and a second movable contact, and the moving direction of the first movable contact and the moving direction of the second movable contact intersect.SELECTED DRAWING: Figure 1

A method for controlling an aircraft comprising at least two engines, each engine being arranged in its own engine compartment, each engine being connected to a fuel supply circuit provided with its own fuel supply cut-off for cutting off a fuel supply, the aircraft comprising at least one fire detector in each engine compartment, the aircraft comprising at least one first extinguisher and one second extinguisher. An automatic assistance phase comprises, in the event of fire being detected in an engine compartment, automatically and with an avionics system, activating the fuel supply cut-off of the engine present in the compartment concerned, then triggering the first extinguisher, then triggering the second extinguisher in predetermined respective conditions.

System (10) for automatically unlocking a door (1) of an aircraft (100) upon impact of the aircraft (100), comprising: a door opening (3); a door (1) which is hinged to the aircraft (100) in the door opening (3); and a locking mechanism (4) designed to lock the door (1) to the aircraft (100) in a closed operating state; wherein the locking mechanism (4) has a load transfer device (5), which is designed and coupled to an impact area (7) of the aircraft (100) such that, in the event of an impact of the aircraft (100) in the impact area (7), the load transfer device (5) transmits a deformation load caused by an inward deformation of the impact area (7) to the locking mechanism (4) and thereby unlocks the door (1).

To provide a flight device capable of guaranteeing safety even when inconvenience occurs to a power connection / disconnection part.SOLUTION: A rotor side power transmission part 26 of a flight device 10 has: a first rotor side power transmission part 261; and a second rotor side power transmission part 262. The power connection / disconnection part 27 transmits power according to a condition and has a first power connection / disconnection part 271 and a second power connection / disconnection part 272. The first power connection / disconnection part 271 is disposed between the first engine side power transmission part 251 and the first rotor side power transmission part 261. The second power connection / disconnection part 272 is disposed between a second engine side power transmission part 252 and the second rotor side power transmission part 262. An operation control part 31 changes a flight mode when the degree to which the first power connection / disconnection part 271 transmits power and the degree to which the second power connection / disconnection part 272 transmits power are apart for a prescribed amount or more.SELECTED DRAWING: Figure 1

An aircraft landing gear assembly (1) is disclosed. The assembly (1) comprises: a leg assembly (1a) comprising: a trailing arm (2) configured to rotate about an axis (3) and carry a wheel (4); a shock absorber (6) coupled to the arm and enabled to dampen rotation of the arm about the axis within a normal operating range, and to limit the rotation of the arm about the first trailing arm axis to the normal operating range; and a fuse member (7) coupled to the shock absorber. The fuse member is configured to fail in the event a vertical load on the wheel exceeds a pre-determined threshold, wherein failure of the fuse member allows the first trailing arm to rotate about the first trailing arm axis beyond the normal operating range.

Disclosed is a crash handling system for an aircraft, comprising an altitude surveillance system adapted to receive altitude data from an altitude sensor and issue an altitude warning signal when a predetermined altitude condition is fulfilled, a crash detection unit, comprising an acceleration detector and issues a crash detection signal when a predetermined acceleration condition is fulfilled, a control unit operatively coupled to the altitude surveillance system and the crash detection unit to receive respective altitude warning signals and crash detection signals; and an energy storage system isolation means operatively coupled to the control unit to electrically isolate an energy storage system of the aircraft upon reception of an activation signal, wherein the control unit transitions from a disabled operation state to an enabled operation state when the altitude warning signal is received; and to issue the activation signal upon reception of the crash detection signal when in enabled state.

The composite sheet comprises a textile component having a first surface and a second surface, and a first polymer coating on the first surface of the textile component, wherein the textile component further comprises at least one woven fabric of continuous filament yarns having an areal weight of from about 70 to about 508 gsm, and a nonwoven fabric comprising a blend of discontinuous fibers of oxidized polyacrylonitrile fibers and silica fibers, wherein the oxidized polyacrylonitrile fibers are present in an amount of from about 30 to about 70 weight percent of the combined weight of the oxidized polyacrylonitrile fibers and the silica fibers, and the nonwoven fabric may be mechanically attached to the woven fabric.

An emergency flotation system for aircraft that, in particular, may be attached to the landing gear or the fuselage of a helicopter, comprising at least two inflatable flotation bodies, each of which are extended in a direction of longitudinal extension and which are spaced apart in a horizontal direction transverse to, in particular, perpendicular to the direction of longitudinal extension, wherein the respective flotation bodies, when inflated, are cross-sectional shaped between their ends that is polygonal perpendicular to their direction of longitudinal extension. An aircraft, in particular, a helicopter, that has an emergency flotation system on its landing gear or fuselage, in particular, wherein the direction of longitudinal extension of the flotation bodies is orientated parallel to the axis of longitudinal extension or transverse, in particular, perpendicular, to the axis of longitudinal extension of the aircraft.

The utility model relates to the technical field of airplane emergency floating, in particular to an emergency floating device for a light airplane. The device comprises a shell, a stop cover and a cover, wing sliding rails are fixed to the two ends of the bottom of the shell, and the stop cover is fixed to one end of each wing sliding rail. According to the utility model, the cover can be quickly unfolded, the disassembly and assembly are very convenient, and the configuration can be quickly adjusted according to task requirements; the shell is installed at the root of a light aircraft wing or the two sides of a helicopter body, streamline design is adopted in the appearance, separation of local airflow is reduced, the installation process is simple and convenient, the internal structure of the aircraft does not need to be transformed, the installation time is greatly shortened, and the installation cost is greatly reduced; the inflation speed of the folding air bag can be obviously increased, the folding air bag can reach the needed pressure in a shorter time, and in addition, the situation that the whole system fails due to the fault of a single air cylinder is avoided.

Provided is a flight device and a flight device control method that ensure safety even if a defect occurs in a power connection / disconnection unit. Rotor-side power transmission units 26 of a flight device 10 include first and second rotor-side power transmission units 261 and 262. Power connection / disconnection units 27 transmit power according to conditions and include first and second power connection / disconnection units 271 and 272. The first power connection / disconnection unit 271 is provided between a first engine-side power transmission unit 251 and the first rotor-side power transmission unit 261. The second power connection / disconnection unit 272 is provided between a second engine-side power transmission unit 252 and the second rotor-side power transmission unit 262. The arithmetic control unit 31 changes the flight status when the degree to which the first power connection / disconnection unit 271 transmits the power and the degree to which the second power connection / disconnection unit 272 transmits the power differ by a certain amount or more.

The invention discloses a data storage and use method of an ejection lifesaving controller, computer equipment and a medium, and relates to the field of ejection lifesaving. The data storage and use method of the ejection lifesaving controller comprises the following steps: electrifying and self-checking the If power-on self-test is normal, whether an ejection signal exists or not is judged, no ejection signal is sent to a receiver, a sensor signal is collected, and periodic self-test is carried out; after the ejection signal is monitored, the authenticity of the ejection signal is judged, and the onboard signal is analyzed; after a differential pressure true value is calculated, the height is determined according to the static pressure true value, the chair-loaded ejection surface speed is calculated, and the true ejection surface speed is determined; determining delay time according to the static pressure real value and the real ejection surface speed, and loading a timer for delaying; after the time delay is finished, a chair-loaded static pressure storage DARM area is collected to determine whether a parachute opening condition is met or not, and a parachute opening instruction is output if the condition is met; and after a parachute opening instruction is output, DRAM data are transferred to an SRAM, and the data of the SRAM are written into an external EEPROM. The data storage and use method of the ejection lifesaving controller can ensure the safety and reliability of data and improve the ejection reliability.