51 results about "Ice protection system" patented technology

An integrated ice detector and ice protection system is provided. The integrated ice detector and ice protection system includes an electric resistance heater element layer, a sensing element, and at least one second electrically insulative layer positioned between the electric resistance heater element layer and the at least one first electrically insulative layer with the sensing element. The sensing element is disposed in at least one first electrically insulative layer and wherein the sensing element is positioned outward from the electric resistance heater element layer.

An assembly for an aircraft propulsion system (20) includes a propulsion module (24), a gas generator (26) and an anti-icing system (28). The propulsion module (24) includes an open propulsor rotor (30). The gas generator (26) is configured to drive rotation of the open propulsor rotor (30) and includes a flowpath (74), a compressor section (69), a combustor section (70), a turbine section (71) and an exhaust section (72). The flowpath (74) extends through the compressor section (69), the combustor section (70), the turbine section (71) and the exhaust section (72). The turbine section (71) is axially between the combustor section (70) and the open propulsor rotor (30) along the axis (22). The anti-icing system (28) is configured to bleed gas from the flowpath (74) in the exhaust section (72) to provide heated gas and to direct the heated gas to the propulsion module (24) to reduce or prevent ice accumulation on an exterior surface (50, 66, 108) of the propulsion module (24).

An assembly is provided for an aircraft propulsion system. This assembly includes a propulsion module, a gas generator and an anti-icing system. The propulsion module includes an open propulsor rotor. The gas generator is configured to drive rotation of the open propulsor rotor about an axis. The gas generator includes a flowpath, a compressor section, a combustor section, a turbine section and an exhaust section. The flowpath extends through the compressor section, the combustor section, the turbine section and the exhaust section. The turbine section is axially between the combustor section and the open propulsor rotor along the axis. The anti-icing system is configured to bleed gas from the flowpath in the exhaust section to provide heated gas. The anti-icing system is configured to direct the heated gas to the propulsion module to reduce or prevent ice accumulation on an exterior surface of the propulsion module.

Aircraft ice protection systems and methods include an aircraft component having an internal layer and an external layer, a piezoelectric actuator element embedded within the aircraft component between material of the internal layer and the external layer, and a control system operably connected to the piezoelectric actuator element and configured to drive actuation of the piezoelectric actuator element to at least one of remove ice formed on the aircraft component and prevent the formation of ice on the aircraft component.

An airfoil of an aircraft with an ice protection system, the airfoil having a leading edge, the ice protection system comprising a first applying mechanism for applying an ice protection fluid along at least a portion of the leading edge of the airfoil, and an ice protection fluid recovering mechanism for recovering the ice protection fluid applied by the first applying mechanism, the ice protection fluid recovering mechanism being arranged downstream of the leading edge of the airfoil. The ice protection system further comprises a second applying mechanism for applying the ice protection fluid recovered by the ice protection fluid recovering mechanism on at least one area of the airfoil arranged downstream of the leading edge of the airfoil. An aircraft may comprise at least one airfoil provided with the ice protection system. A method may be used to protect the airfoil of an aircraft from ice accumulation.

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 electrothermal ice protection system (IPS) installed on an aircraft includes a sensor, a parting strip assembly, and a controller. The sensor monitors a direction of a local incident airflow that is imparted on the sensor. The parting strip assembly is coupled to the critical surface and includes a plurality of heating sections. The controller is in signal communication with the sensor and the parting strip assembly. The controller determines a direction of surface airflow incident on a critical surface of the aircraft based on the local incident airflow and selectively concentrates power to at least one targeted heating section among the plurality of heating sections with respect to non-targeted heating sections among the plurality of heating sections based on the direction of the surface airflow.

The application discloses a kind of tilting rotor aircraft thermal management systems, it is related to aircraft technical field.The system includes: cooling part, wing heat exchange part and cabin environmental control part, each part is communicated by cooling circuit, cooling circuit includes main circuit and branch circuit, main circuit connects cooling part and wing heat exchange part, branch circuit connects cabin environmental control part, main circuit and branch circuit are cooled to the equipment needing cooling using liquid cooling mode, using liquid cooling medium as heat transfer medium, provides cooling control for the equipment needing cooling.The application can realize the interaction and fusion of tilting rotor aircraft airborne cooling system, wing ice protection system, cabin environmental control system, so as to realize the multiple functions of airborne equipment cooling, wing ice protection and cabin temperature management, simplify tilting rotor aircraft equipment type and system, reduce the overall weight.

An ice protection system for use with an aircraft including an electric motor for driving a propulsion structure of the aircraft. The electric motor is cooled by a heat transfer fluid that flows through or by the electric motor to extract heat from the electric motor. The ice protection system uses the heat transfer fluid heated by the electric motor to provide anti-icing and de-icing functionality.

The present application provides a method for controlling the on / off operation of an aircraft engine anti-icing system, the method comprising: determining a critical flight altitude based on specified icing conditions; determining the aircraft's flight Mach number based on the engine operating state being at the maximum inlet Mach number, the engine inlet airflow static temperature being at the icing temperature, and the ambient temperature being at the limit temperature, wherein the aircraft's flight Mach number is the critical flight Mach number; determining the anti-icing system's critical activation temperature based on an anti-icing heat demand formula; activating the anti-icing system when the aircraft's flight altitude, flight Mach number, and engine inlet total temperature are all less than the corresponding critical flight altitude, critical flight Mach number, and anti-icing system's critical activation temperature; and deactivating the anti-icing system when these conditions are less than the corresponding critical flight altitude, critical flight Mach number, and anti-icing system's critical activation temperature. The method provided in the present application, based on flight altitude signals, flight Mach number signals, and engine inlet total temperature signals, provides active anti-icing during the anti-icing system's on / off operation, offering greater safety than passive anti-icing performed after icing using an ice detector.

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.

A hybrid deicing system is disclosed herein. The hybrid deicing system includes a first deicing system coupled to a high critical zone of a control surface, the first deicing system providing thermal energy to the high critical zone of the control surface to remove ice from the control surface, a second deicing system coupled to a low critical zone of the control surface, the second deicing system configured to vibrate the control surface to remove ice from the control surface, and a controller configured to control the first deicing system and the second deicing system.

An electrothermal heater for an ice protection system includes a laminated heater mat which comprises a primary heater element layer, a secondary heater element layer, and at least one dielectric layer interposed between the primary and secondary heater element layers, and control apparatus which comprises power supply apparatus and a current detector. The control apparatus is configured to have (i) a first mode in which the power supply apparatus supplies a heater current to the primary heater element layer and does not supply a heater current to the secondary heater element layer and the current detector monitors the secondary heater element layer for detecting a leakage current of the heater current of the primary heater element layer indicative of burn-out of the primary heater element layer and (ii) a second mode in which the power supply apparatus supplies a heater current to the secondary heater element layer and does not supply a heater current to the primary heater element layer. The control apparatus is configured to switch from the first mode to the second mode in response to detection of said leakage current by the current detector.

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

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.

Aircraft and aircraft ice protection systems include an aircraft environmental control system having a heat exchanger, a supply line fluidly connecting an output of the heat exchanger to a component ice protection element associated with a component of an aircraft, a primary control valve arranged along the supply line, and a controller configured to control operation of the primary control valve to supply air from the heat exchanger along the supply line to the component ice protection element.

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 present invention relates to a method for optimizing a high-temperature warning for wing anti-icing bleed air during flight, belonging to the technical field of aviation wing anti-icing systems. The method comprises the following steps: obtaining the heat exchange efficiency of the aircraft engine bleed air precooler and historical aircraft flight data at different flight altitudes; creating an aircraft wing anti-icing warning temperature expression; calculating the wing anti-icing bleed air high-temperature warning limit based on the anti-icing warning temperature expression and the obtained data; when a high-temperature warning for wing anti-icing bleed air occurs during flight, determining by comparison whether the altitude-temperature at which the wing anti-icing is currently activated is within the calculated high-temperature warning limit for the wing anti-icing bleed air, and taking actions to address the high-temperature warning for the wing anti-icing bleed air based on the optimized judgment result. The present invention can assist pilots in making judgments and effectively reduce the occurrence of abnormal events such as flight return and diversion.

In order to improve the effectiveness and efficiency of ice protection systems, the invention proposes a layered thermo-electric heating member for an ice protection system, preferably for aerial vehicles, the heating member comprising a plurality of active layers and a separation layer between adjacent active layers, wherein each active layer comprises a conductive track that is configured to, when supplied with electrical power, generate a heat power density distribution, wherein each conductive track is configured such that the heat power density distribution is varying depending on a position along at least one active layer dimension of each active layer.

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.

A power switching system can include a main power supply to provide power to electrical devices, such as electric heaters for ice protection during an aircraft operation. During a nominal mode of operation, a subset of electrical devices may be operated sufficiently based on the main power supply. The power switching system may further include an auxiliary power supply to supplement a main power supply for critical phases of operation, such as during takeoff, landing, or flight into known icing conditions. The auxiliary power supply can include a battery or supercapacitor that augments electrical power provided from the main power supply. When an auxiliary mode of operation is needed, such as during the critical phase of flight, an electrical power from an auxiliary power supply can be used to augment the main power supply so that all electrical devices are powered.

The invention discloses an intelligent ship body anti-icing system based on a ship waste heat active intervention mechanism. The intelligent ship body anti-icing system comprises an ice condition sensing device used for obtaining ship body temperature parameters and external environment parameters; the heat energy dispatching and control unit is used for analyzing and processing the parameters acquired by the ice condition sensing device, predicting the ice crystal growth trend and outputting an intervention strategy; the heat guiding assembly is used for connecting a ship waste heat source with a ship body so as to convey waste heat; the heat source interface module is used for adjusting the waste heat conveying amount of the heat guiding assembly; and the main controller is used for receiving the intervention strategy output by the heat energy dispatching and control unit and regulating and controlling the heat guiding assembly and the heat source interface module according to the intervention strategy. Ice condition prediction is achieved through the heat energy dispatching and control unit, intervention is conducted in time before an ice layer is formed, the ship body icing probability is reduced, and the ship body safety is better protected.

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.