68 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.

A target material removal (TMR) element is provided and includes a body region (112); thermal elements (120) within the body region (112); and vibration elements (130) within the body region (112). The TMR element is controlled to perform TMR operations that include one or more of the thermal elements (120) generating heat that is transmitted through the body region (112) 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 (120) beginning generating the heat that is transmitted through the body region (112) to the target material, the one or more of the vibration elements (130) generating vibrations that are transmitted through the body region (112) to the one or more zones of the target material.

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.

A system and a method include a heating layer including perforations extending from a first surface to a second surface opposite from the first surface. The perforations provide acoustic paths. The heating layer is configured to provide electrically resistive heating to a component. An acoustic layer can be coupled to the heating layer. The perforations extend to the acoustic layer.

A system and a method include a heating layer including perforations extending from a first surface to a second surface opposite from the first surface. The perforations provide acoustic paths. The heating layer is configured to provide electrically resistive heating to a component. An acoustic layer can be coupled to the heating layer. The perforations extend to the acoustic layer.

The invention refers to a wind turbine blade formed by two shell halves, wherein one shell half has a first pair of metallic sheets layered, on the outer surface of the shell half, and wherein each of the metallic sheets has been partially embedded within the shell half during the resin infusion process. A first heating mat made of an electrically conductive material is layered on resin-free parts of the metallic sheets, and is also embedded within the shell half so that the first heating mat is electrically in contact with the resin-free parts of the metallic sheets. The invention also refers to a method for manufacturing a wind turbine blade, integrating an Ice Protection System, the method comprising the steps of manufacturing first and second shell halves by means of a resin infusion process, and integrating an Ice Protection System in the same resin infusion process for manufacturing the shell halves.

An ice protection system for an aircraft includes one or more heaters configured to be coupled to a surface of a wing of the aircraft. The ice protection system also includes a controller configured to, in response to a first determination indicating presence of an icing condition, determine a setpoint temperature for a first location of an outer surface of the wing configured to be heated by a heater of the one or more heaters. The controller is also configured to control power provided to the heater based on the setpoint temperature. The setpoint temperature for the first location is determined to cause a second location of the outer surface of the wing, which is heated by the heater and is located at a particular distance from the first location, to remain above a threshold temperature based on a projected temperature profile of the outer surface heated by the heater.

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.

An anti-icing system (200) for a truss reinforced wing (102) of an aircraft (100) is disclosed. The system (200) comprises: a wing (104); a support bar (106); a joint (110) coupling the support bar (106) to the wing (104) to form a structural truss (108) that trusses the wing (102); the magnetic receptor (110) is arranged on the connector (110) or in the connector (110); and an induction heating coil (202) disposed on or within the truss-reinforced wing (102). The induction heating coil (202) is configured to generate an eddy current inside the magnetic susceptor (110) based on driving with an alternating current (AC) excitation current. Eddy current is circulated through the magnetic susceptor (110), thereby causing Joule heating of the magnetic susceptor (110) to prevent icing on the joint (110).

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.

A control system (200) for a de-icing system (400) may comprise an electrical heating mat (402,404), a controller (127), and a tangible, non-transitory memory configured to communicate with the controller, the tangible, non-transitory memory having instructions stored thereon that, in response to execution by the controller, cause the controller to perform operations comprising: receiving (802), by the controller, an enable command from a control interface (175); polling (804), by the controller, a sensor (202) for a sensor data; receiving (806), by the controller, the sensor data from the sensor; passing (808), by the controller, the sensor data to an ice protection logic (204); determining (810), by the controller via the ice protection logic, an icing condition; and applying (812), by the controller, an electrical current to a first busbar portion (604A-1) of the electrical heating mat.

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.

The invention relates to a wing anti-icing bleed air high-temperature early warning optimization method in navigation, and belongs to the technical field of aviation wing anti-icing systems, and the method comprises the steps: obtaining the heat exchange efficiency of an aircraft engine bleed air precooler and the historical flight data of an aircraft at different flight heights; creating an anti-icing early warning temperature expression of the aircraft wing; according to the anti-icing early-warning temperature expression and the obtained data, calculating an anti-icing air-entraining high-temperature early-warning limit of the wing; when wing anti-icing air-entraining high-temperature early warning occurs during navigation, whether the height-temperature for starting wing anti-icing at present is within the calculated wing anti-icing air-entraining high-temperature early warning limit range or not is judged through comparison display, and operation for anti-icing air-entraining high-temperature early warning is adopted according to the optimized judgment result. The method can assist a pilot in judgment, and abnormal events such as flight return and landing preparation are effectively reduced.

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.

The invention 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.

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

An anti-icing system for an aircraft component includes a plurality of heaters. The aircraft component has at least two sections and a joint area. At least one of the heaters is an H-shaped carbon allotrope heater designed to apply heat to the joint area and prevent ice from accumulating in the joint area.

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 refers to a wind turbine blade formed by two shell halves, wherein one shell half has a first pair of me- tallic sheets layered, on the outer surface of the shell half, and wherein each of the metallic sheets has been par- tially embedded within the shell half during the resin infu- sion process. A first heating mat made of an electrically conductive material is layered on resin-free parts of the me- tallic sheets, and is also embedded within the shell half so that the first heating mat is electrically in contact with the resin-free parts of the metallic sheets. The invention also refers to a method for manufacturing a wind turbine blade, integrating an Ice Protection System, the method com- prising the steps of manufacturing first and second shell halves by means of a resin infusion process, and integrating an Ice Protection System in the same resin infusion process for manufacturing the shell halves.