23 results about "Ice protection system" patented technology

A freezing point depressant (FPD) fluid distribution apparatus and ice protection system for a propeller blade, the apparatus comprising an upper section and a lower section, wherein the upper section is configured to distribute FPD fluid to the low pressure side of a propeller blade, and the lower section is configured to distribute FPD fluid to the high pressure side of a propeller blade, and wherein the upper and lower sections each comprise one or more FPD fluid chambers, the or each FPD fluid chamber comprising an associated fluid-permeable cover.

A gas turbine engine including a core air passage, a combustor, and a steam line. The combustor is located in the core air passage and combusts hydrogen fuel producing combustion gases. The steam line is fluidly coupled to the core air passage at a position downstream of the combustor to receive a portion of the combustion gases. A conduit thermally coupled to an external surface of an aircraft may be fluidly coupled to the steam line to receive the combustion gases and to heat the external surface. The gas turbine engine may also include a water vapor condenser fluidly connected to the steam line to receive the combustion gases and to condense the water vapor of the combustion gases. At least one nozzle may be fluidly coupled to the water vapor condenser to inject the condensed water into the core air passage.

A gas turbine engine inlet anti-icing system includes an air supply conduit, a flow control valve, a pneumatic valve actuator, and a mechanical thermostat. The air supply duct has an inlet to receive pressurized air and an outlet coupled to an ice flow prevention duct formed within an inlet portion of the gas turbine engine. A flow control valve is disposed within the air supply conduit between the inlet and the outlet, and is movable to a valve position between a closed position and a plurality of open positions. A pneumatic valve actuator moves the flow control valve between a closed position and a plurality of open positions in response to a pneumatic control pressure. The mechanical thermostat controls the pneumatic control pressure in response to the temperature within the ice flow prevention conduit, thereby controlling the temperature of air in the ice flow prevention conduit.

The present application belongs to the technical field of aircraft electromechanical comprehensive design, and particularly relates to a multi-authorization control system and method of an aircraft air anti-icing system, which comprises an anti-icing switch, an automatic on relay, an automatic off relay, an anti-icing device, an on authorization management device and an off authorization management device, the anti-icing switch is connected with the automatic on relay, the automatic off relay and the anti-icing device are connected in sequence, the on authorization management device is connected with the automatic on relay, and the off authorization management device is connected with the automatic off relay, the multi-authorization control system and method of the anti-icing system are improved according to the existing design of a certain type of helicopter air anti-icing system, and are an anti-icing control system with high automation degree, which saves the energy of pilots and reduces the operation burden of pilots.

An anti-icing system for an exterior surface of an aircraft. The exterior surface is configured to have an airflow over the exterior surface and to have a plurality of zones. At least one heat source is thermally coupled to the exterior surface in each of the plurality of zones. A controller is configured to selectively control the at least one heat source in each of the plurality of zones based on operating conditions related to the airflow over the exterior surface.

Ice protection system for use with an aircraft including an electric motor to drive an aircraft propulsion structure. The electric motor is cooled by a heat transfer fluid that flows through or over the electric motor to extract heat from it. The ice protection system uses the heat transfer fluid heated by the electric motor to provide anti-icing and de-icing functionality. Figure for abbreviation: Fig. 1

An aircraft is provided which includes an airframe, a propulsion system and a thermal anti-icing system. The airframe includes a body and a plurality of wings. The body extends longitudinally along a centerline. The wings are disposed to opposing lateral sides of the body. Each of the wings projects spanwise out from the body. The propulsion system is connected to the airframe. The thermal anti-icing system is configured to heat a region of an exterior surface of the airframe longitudinally upstream of and laterally overlapping with the propulsion system.

The invention belongs to the related technical field of aircrafts, and discloses an efficient anti-icing device and system applied to a low-altitude aircraft and application. The device comprises a group of upper substrates, a group of lower substrates and a thermoelectric layer arranged between the upper substrates and the lower substrates, the thermoelectric layer is an array formed by arranging P-type semiconductors and N-type semiconductors in a staggered and spaced mode, the P-type semiconductors and the N-type semiconductors are connected in series through wires to form a closed loop, and the thermoelectric layer is connected with an external power source to be powered on. The invention further discloses an anti-icing system comprising the deicing device. According to the invention, the problems of low efficiency and high energy consumption in a low-altitude electric aircraft deicing technology are solved.

This invention proposes a method for monitoring early icing of aircraft based on ultrasonic pulse echo. The input parameters of this method include echo signals from ice-free structures, echo signals from iced structures, and the surface temperature of the structures. By analyzing and processing these input signals using the proposed method for early icing warning and ice thickness measurement based on ultrasonic pulse echo, early icing warning information and surface ice thickness data can be obtained, providing a basis for the activation and deactivation of the aircraft icing protection system. This invention can reduce the energy consumption of the aircraft icing protection system when removing ice from the aircraft surface, improving energy efficiency while ensuring flight safety, and has promising application prospects.

The application relates to a distributed electric heating icing protection system and relates to the technical field of electric heating deicing of an airplane, and comprises a surface coating layer covering the surface of the airplane skin, an electric heating layer and a detection layer below the surface coating layer; wherein the electric heating layer and the detection layer are arranged in layers, and are both distributed one-to-one unit structures; the heating components of the electric heating layer of each unit are uniformly arranged; the detection layer of each unit is provided with a detection element for detecting the icing condition; and the electric heating layer and the detection layer of each unit structure are independently controlled by a controller. Compared with the prior art, the uniformity of heat emission can be improved, and the deicing effect is better.

A freezing point depressant (FPD) fluid distribution apparatus and ice protection system for a propeller blade, the apparatus comprising an upper section and a lower section, wherein the upper section is configured to distribute FPD fluid to the low pressure side of a propeller blade, and the lower section is configured to distribute FPD fluid to the high pressure side of a propeller blade, and wherein the upper and lower sections each comprise one or more FPD fluid chambers, the or each FPD fluid chamber comprising an associated fluid-permeable cover.

The invention relates to an aerospace anti-icing technology, in particular to a micro-injection-electric heating composite anti-icing system suitable for a supercooled large water drop condition, which comprises a micro-injection hot air anti-icing module and an electric heating anti-icing module arranged at the rear part of the micro-injection hot air anti-icing module, high-temperature and high-pressure bleed air led out by an engine compressor passes through the flow limiter through the engine bleed air channel, when the wing anti-icing valve is opened, the bleed air flows along the flute-shaped pipe assembly and is sprayed into the micro-injection mixing cavity through the spray holes evenly distributed in the pipe wall of the flute-shaped pipe assembly, peripheral air flow is injected to form efficient hot air flow, and the high-temperature and high-pressure bleed air flows into the micro-injection mixing cavity. A double-skin corrugated plate anti-icing channel is used for enhancing heat transfer and airflow stability. The electric heating anti-icing module comprises an electric heating film embedded into the skin and a temperature sensor. When encountering supercooled large water drop icing meteorological conditions, the device automatically enters the SLD mode, the electrothermal film is started for heating, and the problem that the impact area is increased under the SLD condition is solved. The system significantly improves the anti-icing efficiency, reduces the energy consumption, and has a fault tolerance function.

The invention belongs to the technical field of deicing systems of single-channel branch passenger aircrafts, and discloses a heating control method for valves of a deicing system, which is characterized in that a deicing system controller acquires temperature signals of an airframe temperature sensor, receives static temperature signals from an avionics system DMC, and controls the corresponding valves of the deicing system to heat or cut off heating according to corresponding requirements; and the situation that flight safety is harmed due to abnormal working of the deicing system caused by icing of the valve under the low-temperature condition is prevented.

Embodiments of the present disclosure provide early anti-ice valve fault detection. Engine data associated with a flight operation of an aircraft may be received, the engine data may comprise timeseries data for one or more monitored engine parameters and the aircraft may be associated with an anti-ice system comprising a thermal anti-ice valve and a pressure sensor. One or more feature datasets may be extracted from the engine data and using one or more feature extraction models. Each feature dataset may represent a data slice from the engine data that satisfies thermal anti-ice valve feature extraction criteria. An anti-ice valve fault prediction may be generated by applying the feature dataset to one or more fault prediction models. The performance of one or more prediction-based actions may be initiated based on the anti-ice valve fault prediction.

The invention relates to a distributed electric heating icing protection system, which relates to the technical field of aircraft electric heating deicing, and comprises a surface coating layer covering the surface of aircraft skin, and an electric heating layer and a detection layer below the surface coating layer, wherein the electric heating layer and the detection layer are arranged in a layered mode and are each of a distributed one-to-one unit structure, heating assemblies of the electric heating layer of each unit are evenly arranged, the detection layer of each unit is provided with a detection element used for detecting the icing condition, and the electric heating layer and the detection layer of each unit structure are independently controlled by a controller. Compared with the prior art, heat dissipation uniformity can be improved, and the deicing effect is better.

A target material removal (TMR) element is provided and includes a body region; thermal elements within the body region; and vibration elements within the body region. The TMR element is controlled to perform TMR operations that include one or more of the thermal elements generating heat that is transmitted through the body region to a target material in a manner that defines one or more zones of the target material. The TMR operations further include, subsequent to the one or more of the thermal elements beginning generating the heat that is transmitted through the body region to the target material, the one or more of the vibration elements generating vibrations that are transmitted through the body region to the one or more zones of the target material.

The application relates to the technical field of cables, in particular to an overhead conductor, a preparation method of the overhead conductor and an overhead conductor anti-icing system. The overhead conductor comprises a core body and a stranded layer stranded on the periphery of the core body; the core body is in a hollow structure, and the hollow structure of the core body forms a first hot air passage; and the outer surface of the stranded layer has a turbulence structure. The overhead conductor is cooperated with the active ice-melting structure and the passive turbulence structure, synergistically, the anti-icing performance of the overhead conductor is greatly improved, and the electric loss is small.

The present application belongs to the field of helicopter inlet anti-icing design, and relates to a flow-adjustable helicopter inlet hot air anti-icing system, which comprises an inlet skin, a heat preservation skin, multiple partition frames, a flow control valve, an atmospheric temperature sensor, a control device and a gas supply pipeline. The flow of hot air into the anti-icing cavity can be automatically adjusted according to the ambient temperature, and heat waste can be avoided under the condition of meeting the anti-icing demand. The hot air in the inlet anti-icing cavity flows from back to front, which can effectively reduce the influence on the engine inlet air distortion. The multiple partition frames are used to distribute the hot air in the anti-icing cavity, which can reduce the height of the anti-icing cavity to 5-15 mm and improve the heat exchange efficiency. The heat preservation layer wrapped on the inlet heat preservation skin can effectively reduce heat loss.

A system for controlling ice accumulation on a surface of an aircraft, the system includes a carbon nano-tube (CNT) heater comprising: a CNT layer; a first encapsulation layer disposed on a first side of the CNT layer formed of a first encapsulation layer thermoplastic material; and a second encapsulation layer disposed on a second side of the CNT layer formed of a second encapsulation layer thermoplastic material. The system also includes a fore composite structure that includes a fore composite structure thermoplastic material disposed on the first side of CNT heater, an aft composite structure that includes an aft composite structure thermoplastic material disposed on the first side of CNT heater and a sensor layer disposed between the CNT heater and the one of the fore and aft composite structures.

An aircraft wing may comprise an airfoil having deicing zone, an anti-icing zone, and an ice runback control zone. An aircraft wing may comprise an electro-thermal ice protection system disposed in the aircraft wing. The electro-thermal ice protection system may be disposed along the deicing, anti-icing, and ice runback control zones of the airfoil to improve aerodynamic performance of the aircraft and reduce ice formation along the wings of the aircraft.

The utility model provides an anti-icing system of an aircraft windshield, relates to the technical field of aircraft electric environment control, and is used for improving the anti-icing efficiency of the aircraft windshield. The system comprises a one-way valve, an airflow rectifier and an air film slit, an air inlet of the air film slit is located in a cockpit, and an air outlet of the air film slit is located at the bottom of an aircraft windshield. The airflow rectifier and the one-way valve are located in a cockpit, one end of the airflow rectifier is connected with an air inlet of the air film slit, and the other end of the airflow rectifier is connected with the one-way valve; the airflow rectifier is used for reducing the turbulence degree of airflow flowing in through the one-way valve.