515results about "De-icing equipments" patented technology

The application relates to the field of aviation deicing technology, and particularly discloses a low-energy-consumption deicing system and control method for limited onboard energy, which comprises: a plurality of deicing subzones, which are arranged symmetrically along an aircraft body axis; each deicing subzone is provided with independent heating devices and vibration devices; the heating devices are used for melting ice layers to reduce the adhesion between the ice layers and the skin; the vibration devices are used for making the skin produce pulse vibration so that the ice layers on the skin fall off; and a subzone control module is used for carrying out subzone round control on the plurality of deicing subzones, so that the plurality of deicing subzones carry out deicing work in turns, and the deicing subzones carrying out deicing at the same time are symmetric along the aircraft body axis. The application greatly reduces the energy demand of the deicing technology while ensuring the deicing efficiency.

A gas turbine engine is provided that includes a fan section, a compressor section, a combustion section, a turbine section, an annular fan bypass air path, a bleed air structure, an annular compartment, and an anti-icing system. The compressor section has a low and high pressure compressor sections. The fan bypass air path is defined by inner and outer bypass wall structures. An outer radial boundary structure defines an outer radial boundary of a core gas path within the low pressure compressor section. The bleed air structure is disposable in an open state or in a closed state. The annular compartment is defined by the inner bypass wall structure, the outer radial boundary structure, and a firewall. The anti-icing system is disposable in an activated state or in a deactivated state. In the activated state the anti-icing system is configured to mitigate ice accumulation within the annular compartment.

The invention belongs to the technical field of electric heating deicing, and particularly relates to a surface heating electric heating film and a preparation method thereof. Comprising an outer surface insulating layer, a heating layer, a power supply electrode and an inner surface insulating layer, the heating layer is prepared from a carbon-based conductor composite polymeric material in a printing or coating mode, and the carbon-based conductor composite polymeric material is formed by mixing polyimide resin and carbon fibers (or one or more of graphene, carbon nanotubes and the like). The surface heating electric heating film provided by the invention has the characteristics of simple and reliable structure, light weight and surface heating performance, can be applied to ice prevention and removal operation of parts such as aircraft wings, empennages, propellers and the like, and has relatively high practical application value.

A system and a method include a control unit for determining a probable de-icing status for one or more aircraft scheduled to depart from one or more airports. The control unit is configured to determine the probable de-icing status based on a current date at the one or more airports, current weather at the one or more airports, and the actual de-icing status of a plurality of prior aircraft scheduled to depart from the one or more airports before the one or more aircraft.

The invention relates to the technical field of aircraft electric heating deicing, in particular to a pulse heating aircraft deicing method and system.The method comprises the following steps that the ice layer thickness, the ice layer type, the airspeed and the attack angle of each aircraft surface area are collected, the inherent pneumatic sensitivity coefficient of each area is called, the aerodynamic penalty index of each area is established, and the aerodynamic penalty index of each area is calculated; and according to the values of the aerodynamic penalty indexes of the areas, performing descending order arrangement to generate a deicing area priority sequence. Through multi-parameter joint evaluation of the ice layer thickness, the ice layer type, the airspeed and the incidence angle, the heating resource distribution sequence is optimized, energy waste is avoided, in the local heating execution stage, temperature change slope real-time monitoring is adopted to replace fixed time control, the actual thermal response behavior of the ice layer can be dynamically responded, and the real-time thermal response performance of the ice layer is improved. The initial phase change process of the ice crystal is identified through a phase change starting slope threshold value, and then a medium-power pulse control mode is switched, so that a steady-state regulation heating path is formed, and the phase change control continuity is strengthened.

The invention discloses an airplane icing early warning method, and relates to the technical field of icing early warning. According to the airplane icing early warning method, cloud picture data and cloud product file data of a set airplane are obtained, a cloud field operation state feature set of the set airplane is extracted based on the cloud product data of the set airplane, and a cloud liquid water path value of the set airplane is analyzed; based on a pre-trained cloud picture error mapping model and in combination with cloud picture data of the set airplane, analyzing a cloud type temperature correction factor of the set airplane, and in combination with the cloud field operation state feature set, analyzing an optimized cloud top temperature value of the set airplane; based on the cloud liquid water path value of the set airplane, the cloud top temperature value is optimized, and the ice accretion index of the set airplane is analyzed, ice accretion early warning processing is performed on the set airplane based on the ice accretion index, so that the finally generated ice accretion index result has higher environmental adaptability, dynamic correction and high-precision characterization of the ice accretion risk are realized, and the method is suitable for large-scale popularization and application. And thus, the accuracy of ice accumulation early warning is improved.

The invention discloses a laser processing microstructure electric heating anti-icing and de-icing skin and a using method thereof.The laser processing microstructure electric heating anti-icing and de-icing skin comprises a skin body arranged on the surface of an airplane, the skin body comprises an electric heating microstructure anti-icing and de-icing skin body, a microstructure anti-icing and de-icing skin body and a common skin body, and the input end of the electric heating microstructure anti-icing and de-icing skin body is connected with an airborne power source through a wire; the electric heating micro-structure anti-icing and deicing skin is arranged on the windward side, with the maximum airflow collection coefficient, of the outer surface of the aircraft, the micro-structure anti-icing and deicing skin is arranged on the upstream and downstream of the windward side of the outer surface of the aircraft, and the common skin is arranged on other parts, except the electric heating micro-structure anti-icing and deicing skin and the micro-structure anti-icing and deicing skin, of the outer surface of the aircraft. According to the laser processing microstructure electric heating anti-icing and deicing skin and the using method thereof, through matched arrangement of the electric heating microstructure anti-icing and deicing skin and the microstructure anti-icing and deicing skin, the purposes of being good in anti-icing and deicing effect, low in energy consumption, light in weight and good in maintenance performance can be achieved.

The invention relates to a method for detecting aircraft icing of an aircraft, the method comprising the following steps: - Recording at least a part of an outer flow surface of the aircraft by means of at least one optical image sensor of a detection device and generating digital image data containing the recorded part of the outer flow surface of the aircraft, - inputting the digital image data into a machine learning system of the detection device as input data, which has learned a correlation between digital image data as input data and various types of aircraft icing as output data by means of at least one machine-learned decision algorithm in order to obtain output data, and - Detecting aircraft icing depending on the output data received and, if aircraft icing has been detected, classifying the aircraft icing with regard to at least one icing type depending on the output data received.

The invention relates to a system (10) for de-icing the blades (14) of the propeller (12) of an aircraft, the system comprising an H-bridge inverter device (40); a rotary transformer (22) comprising a main primary winding (26), a primary measurement winding (28), and a secondary circuit (30) delivering a secondary alternating voltage (Vs); a rectifier device (50); a de-icing device (20); a control device (60) configured to control the switching elements of the inverter device in an out-of-phase manner, using the value of the voltage (Vpm) across the terminals of the primary measurement winding, measured during a measurement interval of predetermined duration starting at the instant (t11,t33) when the inverter device switches from a controlled state to a short-circuit state.

The invention relates to the technical field of aviation equipment, and discloses a wing-shaped structure suitable for anti-icing and deicing, which comprises a shell and an electric heating piece, an electric heating piece is arranged in a shell, a first heating layer is formed on the upper wing face and the lower wing face of the inner wall of the shell, a second heating layer is formed on the front edge of the inner wall of the shell, the first heating layer and the second heating layer are electrically connected, and the first heating layer and the second heating layer are connected with an external power source through an electrode layer so that power can be supplied to the electric heating piece, and temperature distribution of different parts can be controlled. After the electrode layer conducts electricity, a first temperature area is formed on the first heating layer, a second temperature area is formed on the second heating layer, and the temperature of the first temperature area is lower than that of the second temperature area, so that it is guaranteed that the front edge of the airfoil structure is kept at the high temperature, and the deicing and anti-icing requirements are met. According to the airfoil structure suitable for anti-icing and deicing, differentiated heating can be conducted according to icing characteristics of different parts of the airfoil structure, it is guaranteed that the anti-icing and deicing requirements are met, meanwhile, energy consumption is reduced, and the energy utilization rate is increased.

The invention provides an electric heating shock-resistant leading edge and a using method thereof.The electric heating shock-resistant leading edge comprises an outer side skin, a leading edge heating layer, a leading edge inner skin, an energy absorption structure layer, an energy absorption inner skin, ribs and a control module which are sequentially arranged in a stacked mode from outside to inside, the energy absorption structure comprises a Z-shaped body and a supporting body, one end of the supporting body is connected with the end of the Z-shaped body, the other end of the supporting body is connected with a bent portion of the Z-shaped body, the Z-shaped body is connected to the front edge inner skin, the supporting body is made of negative thermal expansion materials, and a probe is arranged at the position, corresponding to the supporting body, of the rib and comprises a pressure sensor and a distance monitor. And the probe and the front edge heating layer are electrically connected to the control module. The electric heating deicing function of bird impact resistance and automatic control of the front edge structure is achieved, the temperature control reliability is improved, the structural weight is reduced, and the technical problems of temperature control, bird impact resistance, structural weight reduction and the like are solved.

The application discloses a kind of deicing system real-time state evaluation and fault prediction method and system, the method includes: through the ground state monitoring test of deicing system simulates deicing system fault state, obtains the sensor data under different fault states;Different fault states are used to train state evaluation model and fault prediction model by the sensor data set constructed with data;Sensor is installed on the aircraft deployed with deicing system, to collect the working data of deicing system;The data collected by sensor is input into trained state evaluation model and fault prediction model, the real-time state of deicing system is evaluated, and the possible fault is predicted.The application can evaluate current anti-icing system working state according to current state monitoring data, and predict the possible fault of system.

Provided are embodiments for method and system for performing an integrated ice protection with prognostics and health management using fiber optic sensors. Embodiments can include reading a signal from each sensor of an array of sensors installed on a surface of a structure or equipment, wherein each sensor is a fiber optic sensor, and generating a map based on reading the signal from each sensor, wherein the map monitors a condition of the surface detected by each sensor. Embodiments can also include determining at least one of an abnormal condition or a failure based at least in part on reading the signal from each sensor; and performing at least one of adjusting power control for the structure or equipment or communicating the abnormal condition or failure of the structure or equipment.

The air inlet functional lip component of the unmanned aerial vehicle comprises a functional lip, the outer end of a functional lip body is provided with an outer side connecting part, and the inner end of the functional lip body is provided with an inner side connecting part; the outer side connecting part comprises an elastic shifting piece fixed to the outer end of the functional lip body, the free end of the elastic shifting piece is provided with a locking block, and the locking block is matched with a locking groove formed in the outer wall of the air inlet channel; the inner side connecting part comprises a positioning pin, and the positioning pin is matched with a positioning sleeve on the inner wall of the air inlet channel; the locking block on the elastic shifting piece is pressed to the inner side of the outer wall of the air inlet channel, moves forwards and then is inserted into the locking groove to be locked, so that the positioning pin is inserted into the positioning sleeve while the functional lip body and the outer wall of the air inlet channel are fixed, and the functional lip body and the inner wall of the air inlet channel are fixed. The component can realize quick disassembly and assembly, connection stability and sealing performance of the lip, and the maintenance efficiency and reliability of the unmanned aerial vehicle are improved. The invention further discloses an unmanned aerial vehicle which has the same advantages.

An aircraft propulsion system includes aircraft systems having at least one hydrogen tank and an aircraft-systems heat exchanger and engine systems having at least a main engine core, a high pressure pump, a hydrogen-air heat exchanger, and a turbo expander (800). The main engine core includes a compressor section, a combustor section having a burner, and a turbine section. Hydrogen is supplied from the at least one hydrogen tank through a hydrogen flow path, passing through the aircraft-systems heat exchanger, the high pressure pump, the hydrogen-air heat exchanger, and the turbo expander (800), prior to being injected into the burner for combustion. The turbo expander (800) includes a rotor (804) separated into a first expander portion (812) and a second expander portion (814) arranged about an output shaft (806) and the output shaft (806) is operably connected to a generator configured to generate electrical power.

The invention relates to an anti-icing mat (1) for an aircraft part (P), the mat (1) comprising: at least one contact layer (10) for application to a surface (S) of said part (P), an encapsulation layer (15), and a resistive structure (20) interposed between the contact layer (10) and the encapsulation layer (15), the resistive structure (20) having at least one heating layer (22) capable of generating a heat flux through the surface (S), made of a material comprising a nonwoven of carbon nanotubes impregnated with an epoxy resin, the heating layer (22) being in the form of a pattern comprising a plurality of resistive segments (23) electrically insulated from each other by the encapsulation layer (15). Figure for the abstract: Fig. 4

This invention discloses an aircraft icing early warning method, relating to the field of icing early warning technology. The method acquires cloud image data and cloud product file data for a specified aircraft. Based on the cloud product data, it extracts the cloud field operation status feature set of the specified aircraft and analyzes the cloud liquid water path value. Based on a pre-trained cloud image error mapping model and combined with the cloud image data of the specified aircraft, it analyzes the cloud shape temperature correction factor of the specified aircraft and, combined with the cloud field operation status feature set, analyzes the optimized cloud top temperature value of the specified aircraft. Based on the cloud liquid water path value and the optimized cloud top temperature value, it analyzes the icing index of the specified aircraft. This invention performs icing early warning processing on the specified aircraft based on the icing index, thereby making the final generated icing index result more environmentally adaptable and achieving dynamic correction and high-precision characterization of icing risk, thus improving the accuracy of icing early warning.

Devices, systems, and methods for remote deicing / anti-icing of an aircraft are described herein. One remote operator deicing / anti-icing station located in a remote location from a deicing / anti-icing rig includes a number of remote rig control components for actuation by a remote operator, wherein the remote rig control components include a remote operator user interface, located outside of a rig being at least partially controlled by the remote operator deicing / anti-icing station, the rig including a truck and a deicing / anti-icing control cabin that controls disbursement of substances to deice / anti-ice an aircraft, and a communication device that communicates instructions from the remote operator user interface and the number of remote rig control components to the rig to control at least one of the truck and the deicing / anti-icing control cabin.

An inlet cone (20) for an aircraft turbine engine (1) is configured to be rotated about an axis (X). The inlet cone comprises a frustoconical body (22) and a tip (21) made of elastically deformable material fixed to a smaller-diameter end (221) of said body. The body comprises through-housings (222) distributed around the axis, and parts (23) made of elastically deformable material located in the through-housings. Each of the parts (23) made of elastically deformable material comprises at least one peripheral flange (230) for retaining the part in the corresponding through-housing (222). The peripheral flange bears against an internal surface (22a) of the body.

The invention discloses an unmanned aerial vehicle anti-icing and de-icing comprehensive logic control method and system, and the method comprises the steps: enabling an airborne control computer to judge whether a precondition for starting unmanned aerial vehicle anti-icing and de-icing is met or not according to the received power supply data sent by a main power supply and the unmanned aerial vehicle power data sent by a piston engine; the airborne control computer receives data sent by a sensor carried by the unmanned aerial vehicle, and judges whether the necessity of starting the unmanned aerial vehicle to prevent and remove ice is met or not; secondarily determining whether it is necessary to start the unmanned aerial vehicle to prevent and remove ice; determining an anti-icing and de-icing operation mode of the unmanned aerial vehicle; the airborne control computer issues an instruction to start the unmanned aerial vehicle for ice prevention and removal, and an ice prevention and removal execution device receives the instruction to complete corresponding actions; the airborne control computer collects the anti-icing and deicing operation data of the unmanned aerial vehicle, monitors the operation state, and automatically isolates and processes faults occurring in the operation process. According to the invention, unmanned aerial vehicle ice prevention and removal can be fully started as required, and unmanned aerial vehicle energy is saved.

A jet aircraft having an airflow modification device in its engine duct, the airflow modification device comprising an inlet, an outlet, and a flow path extending from the inlet to the outlet, wherein the device is positioned to receive an incoming airflow and is configured to at least partially modify the airflow within the flow path to improve the ordering of the incoming airflow, the airflow modification device further comprising at least one pitot sensor located at the inlet and integrated with the device, and a pressure line having a first end and a second end, the first end being connected to the pitot sensor and the second end being connected to a transducer, wherein the pressure line is completely enclosed within the device.

Aspects of the present disclosure can be embodied in a systems for heating an exterior surface of a bulk medium. In one example, the system includes two or more coupling strips spaced apart from one another and attached to the bulk medium. Each of the coupling strips has a multi-layer structure extending along a surface of the bulk medium that forms, in combination with the bulk medium, an electrical transmission line. The multi-layer structure includes a first dielectric layer over the bulk medium, a conductive layer over the first dielectric layer, a second dielectric layer over the conductive layer, and a conductive shielding layer over the second dielectric layer. A power control system is coupled to the conductive layer of each of the coupling strips and to the bulk medium. The power control system is configured to heat the bulk medium by providing current to the coupling strips.

The invention belongs to the technical field of composite material weaving, and particularly relates to an electric heating material woven tubular structure, electric heating material woven cloth and an electric heating layer structure. The braided structure is composed of an outer-layer glass fiber tubular structure and an inner-layer carbon nano tube conductive wire, the carbon nano tube conductive wire is wrapped in the glass fiber tubular structure in a spiral winding mode, and a composite material braided bundle which is high in strength, insulated, light in weight and conductive is formed. The two-dimensional woven bundle is further made into woven cloth, the surface heating function is achieved through the good conductivity of the woven cloth, and the method is particularly suitable for heating assembly arrangement of structural parts such as complex curved surfaces or airfoils. The technology overcomes the defects of traditional heating materials in the aspects of shape adaptability and strength, and has wide application prospects.

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.

The invention relates to a heated leading edge structure for an aircraft, a slat, a wing and an aircraft. A heated leading edge structure (14) for an aircraft (10) includes a leading edge panel (20) having an outer surface (32) configured for contact with an ambient flow (A) and an inner surface (34) opposite the outer surface (32); a reinforcing member (22) connected to the inner surface (34) of the leading edge panel (20); and a heating device (24) configured to heat the leading edge panel (20); the heating device (24) comprises at least one indirect heating device (44) mounted on the reinforcing member (22), said indirect heating device (44) being configured for heating the leading edge panel (20) at least by means of conduction through the reinforcing member (22).