221 results about "Ice protection system" patented technology

An anti-icing system for an aircraft engine nacelle comprises a generally ring-shaped hollow spray tube for directing hot gasses toward a portion of the nacelle; a plurality of fasteners for attaching the spray tube to a support structure of the aircraft engine; and a supply duct for delivering the hot gasses from the aircraft engine to the spray tube. The spray tube includes two closed ends which define a thermal expansion gap therebetween to accommodate thermal expansions and contractions caused by the hot gasses. Each of the fasteners comprises a support bracket for attachment to the support structure of the aircraft and a guide assembly for supporting the spray tube to the support bracket. The guide assembly includes a generally ring-shaped block for encircling the spray tube, a bushing positioned within the block for contact with a bearing sleeve of the spray tube, and a pair of bushing retainers positioned on opposite sides of the block for retaining the bushing within the block. The bushings permit the spray tube to slip relative to the fasteners to further accommodate thermal expansions and contractions.

A system and method for ice protection of a component, wherein the system is adapted to be adhesively bonded to a surface of the component. The system includes a heating element layer, at least one thermally conductive adhesive layer that adhesively bonds a first side of the heating element layer to the component, an insulation layer, at least one thermally insulating adhesive layer that adhesively bonds a second side of the heating element to the insulation layer, an electrical bus bar adapted to provide a connection between a power supply and the heating element layer, and at least one temperature sensor.

An ice protection system for de-icing an aircraft and / or for keeping the aircraft free of ice includes a composite structure for an aircraft component and a structural element of an aircraft with an ice protection system that is integrated in the composite structure. The ice protection system includes a flat surface-like electric heating element that is embedded in the composite structure below an outer layer of an outer surface of the aircraft component.

The invention relates to an anti-icing system of an inlet lip of an engine. The system comprises an air guiding tube, a pressure adjusting shut-off valve, a first temperature sensor, a second first temperature sensor and a controller, wherein an air inlet of the air guiding tube is connected on a high-pressure compressor to lead-in hot air of the high-pressure compressor while an air outlet of the air guiding tube is formed in a circular chamber of the inlet lip to release the hot air; the pressure adjusting shut-off valve is distributed on the air guiding tube and is used for adjusting the flow of the hot air in the air guiding tube; the first temperature sensor is distributed in a predetermined hottest region of the inlet lip of the engine; the second temperature sensor is distributed in a predetermined coldest region of the inlet lip of the engine; and the controller is electrically connected with the first temperature sensor, the second temperature sensor and the pressure adjusting shut-off valve and can be used for adjusting the openness of the pressure adjusting shut-off valve according to the temperature T1 sensed by the first temperature sensor and the temperature T2 sensed by the second temperature sensor, so that a covering of the inlet lip is not damaged and is not frozen by the temperature of the inlet lip.

A nacelle inlet lip (22) for an aircraft engine (16) comprises a structural body (30) and an ice protection system (60). The ice protection system (60) comprises a plurality of ice-protectors (62, 64, 66) independently controllable to allow inner zones (52, 54, 56) to be operated in different modes, different power levels, different time intervals and / or different energy amounts. In this manner, an inner aft zone (52), an inner mid zone (54), and an inner aft zone (56) can be anti-iced and / or de-iced to best preserve desired airflow patterns, to minimize ice particle shed size, and to optimize power consumption.

An icing protection system and method for enhancing heat transfer includes a substrate having an inner wall, an outer wall and a thickness separating the inner wall and the outer wall. A metallic layer deposited on the inner wall of the substrate by an electric arc thermal spray deposition process using at least one metallic wire has a thickness between about 0.203 mm (0.008 inches) and about 0.432 mm (0.017 inches), a surface roughness greater than about 12.7 microns (500 micro-inches) Ra, and a heat transfer augmentation of at least about 1.1. The metallic layer is formed on the inner wall from an M-Cr—Al alloy where M is selected from Fe, Co and Ni. The metallic layer defines a plurality of turbulators that act as micro-fins to enhance heat transfer from a heated gas in flow communication with the metallic layer through the substrate to prevent the formation of ice on the outer wall.

The ice prevention system of the present invention utilizes, in combination, a de-ice subsystem and anti-ice subsystem in a new configuration on an aircraft with engines located behind its wings. Depending on the vertical location of the engines with respect to the wings, the anti-ice subsystem is configured on an upper or lower portion of a leading edge of the wings. If the engine is located above the wing, the anti-ice system is configured only on a section of the wing aligned with the engine and only on the upper portion of the leading edge. If the engine is located below the wing, the placement of anti-ice system is reversed. The de-ice subsystem is configured on the opposite portion of the leading edge adjacent to the anti-ice subsystem and on the remainder of the leading edge of the wing.

An electro-thermal ice protection system for an airfoil comprises: a metal foil heater including an integral parting strip, the metal foil heater configurable to cover at least a portion of a leading edge of the airfoil with the integral parting strip disposed along an air-stagnation zone of the leading edge; and a controller coupled electrically to the metal foil heater for controlling electrical energy from a power source to the metal foil of the heater in accordance with a pulse duty-cycle and for controlling power to the parting strip of the heater to maintain the air-stagnation zone virtually free of ice formation. A method of making an exemplary metal foil heater for use in ice protection of an airfoil comprises the steps of: preparing a sheet of Titanium foil in a rectangular shape having a length substantially greater than a width thereof for covering at least a portion of the leading edge of the airfoil; applying a layer of Nickel material over an area along each width edge of the Titanium foil; applying a layer of Copper material over the Nickel layer along each width edge of the Titanium foil; and attaching conductor wires to the Copper layers at each width edge.

An aircraft engine nacelle inlet is provided with an inlet cowling. The inlet cowling includes an inner lip, an outer lip, and a leading edge portion connecting the inner and outer lips. Heating elements are provided proximate the leading edge, either on an inside surface of the cowling or on an outside surface. An inner barrel portion and an outer barrel portion of the nacelle inlet define a space therebetween. Ice protection-related equipment such as controllers, cables, switches, connectors, and the like, may reside in this space. One or more access openings are formed in the outer barrel to enable an operator to gain access to this equipment. The inlet cowling attaches to the inner and outer barrels with its outer lip extending sufficiently far in the aft direction to cover the access opening. When the cowling is removed, the access opening is uncovered, thereby permitting access to the equipment.

An anti-icing system and method for an aircraft engine nacelle. The nacelle has an inlet lip that defines a leading edge of the nacelle and has a cross-sectional shape and oppositely-disposed exterior and interior surfaces. An anti-icing system contacts at least the inlet lip of the nacelle. The anti-icing system includes at least one heating element having a cross-sectional shape that conforms to the cross-sectional shape of the inlet lip and in which carbon nanotubes are oriented and arranged to conduct electrical current through the heating element. Passing an electrical current through the heating element causes Joule heating of the heating element and heating of the inlet lip by thermal conduction.

An aircraft, an aircraft engine and a corresponding method provide for anti-icing based upon energy provided by a fan assembly of the aircraft engine, such as by rotation of a low pressure shaft of the fan assembly. The aircraft includes an aircraft body having wings and an aircraft engine carried by the aircraft body. The aircraft engine includes a core gas turbine engine and a fan assembly coupled to the core gas turbine engine. The fan assembly may include a fan, a low pressure turbine and a low pressure shaft connecting the fan and the low pressure turbine. The aircraft engine also includes an electric anti-icing system that has a variable frequency generator driven by the fan assembly, and one or more resistive heaters carried by the wings. The variable frequency generator is configured to provide electrical energy to the one or more resistive heaters.

The invention discloses a wing test element for an ice wind tunnel of an anti-icing system for an aircraft wing. The wing test element comprises a shell, an anti-icing cavity and a test flute pipe, wherein the shell is formed by performing sweepback stretching on a critical cross-section; the critical cross-section is defined in the following mode: a possible serious icing area is determined according to the water droplet impingement property of an external flow field of the aircraft wing, the most possible serious icing area is determined according to thermal power distribution of a flute pipe in the aircraft wing, and a chordwise airfoil along the airflow in the most possible serious icing area is defined as the critical cross-section; the sweepback stretching is that the critical cross-section is stretched into a test wing and the aircraft wing leading edge wire in the wing unfolding direction of the test wing and a straight line vertical to the critical cross-section form a sweepback angle; the anti-icing cavity is formed in the shell, and structural parameters of the anti-icing cavity are configured to be the same as those of an anti-icing cavity of a true aircraft wing at a standard position; and the test flute pipe is arranged in the anti-icing cavity, and related parameters of the test flute pipe are configured to be the same as parameters of a flute pipe of the true aircraft wing at the standard position.

A modular fan inlet ice protection system includes a center ring, an outer ring and a plurality of struts. The struts are each coupled at a first end to the center ring, and at a second end to the outer ring. A plurality of shell members are each configured to be removably coupled to a corresponding strut. Moreover, the shell members each include an electrothermal heater element Preferably the system further includes a plurality of inlet guide vanes each including an electrothermal heater element.

A leading edge component (30) for an aircraft. The component includes at least one supporting rib (40); a skin of the leading edge; and an ice-protection system comprising a thermoelectric heater mat (10) located between the rib and the skin, the heater mat comprising an electrically resistive heater element (12, 22) embedded in a thermoplastic layer (16, 26, 28), wherein the rib is bonded to the skin by the thermoplastic of the heater mat. A method making a leading edge component for an aircraft. The method includes: providing at least one rib, a skin of the leading edge, and an ice-protection system comprising a thermoelectric heater mat having an electrically resistive heater element embedded in a thermoplastic layer; locating the heater mat between the rib and the skin; and bonding the rib to the skin using the thermoplastic of the heater mat.

An icing protection system comprises a fuel cell, an evaporative cooling system, a thermal deicing device, a control unit and a device for measuring the pressure in the evaporative cooling system. The evaporative cooling system is in thermal contact with the fuel cell and with the thermal deicing device in order to conduct heat from the fuel cell to the thermal deicing device. The control unit controls the electrical power demanded from the evaporatively cooled fuel cell in dependence on a deviation of the pressure measured by the pressure measuring device from a predetermined NOMINAL pressure. In this way, an energetically very efficient and very well controllable icing protection system is made available that can be operated without bleed air supplied by engines or electrical energy provided by engine generators.

An electrothermal heater mat (3) is provided for an ice protection system for an aircraft (1) or the like. The heater mat (3) is a laminated heater mat and comprises dielectric layers (50-58), a heater element (501) and a conductive ground plane (71-74) for detecting a fault with the heater element (501). The dielectric layers (50-58) are made of thermoplastic material, and the ground plane (71-74) is formed by spraying metal such as copper onto the thermoplastic material of one of the dielectric layers.

An ice protection system and method are provided that facilitate the prevention or reduction in ice formation upon an aerodynamic surface. The ice protection system may include at least one induction coil spaced apart from an electrically conductive layer of an aerodynamic surface and disposed interior of an exterior surface of the aerodynamic surface. For example, the ice protection system may include a plurality of induction coils with each induction coil being spaced apart from the other induction coils such that the respective induction coils are associated with different portions of the aerodynamic surface. The ice protection system may also include a signal source configured to provide alternating current to the induction coil(s) to thereby generate an electromagnetic field that creates an eddy current in the electrically conductive layer so as to heat the electrically conductive layer.

A system and method for ice protection of a component, wherein the system is adapted to be adhesively bonded to a surface of the component. The system includes a heating element layer, at least one thermally conductive adhesive layer that adhesively bonds a first side of the heating element layer to the component, an insulation layer, at least one thermally insulating adhesive layer that adhesively bonds a second side of the heating element to the insulation layer, an electrical bus bar adapted to provide a connection between a power supply and the heating element layer, and at least one temperature sensor.

An ice detection system, in particular to a system for detecting icing conditions and a system for detecting the presence of ice formed on a structure, and the methods for doing the same. Two heaters of the ice protection system (experiencing substantially the same environmental conditions) are driven to different temperatures (the first greater than or equal to 0° C., and the second less than 0° C.). The difference in the powers required to drive the two heaters indicates the prevailing environmental conditions. A heater of the ice protection system is driven and the rate of change of surface temperature is measured over several periods. A substantial deviation from a rate of change of surface temperature, which indicates the presence of ice on the surface of the structure, is detected.

Aircraft air scoop icing protection systems and methods include an air duct having air inlet and outlet openings and a fairing which covers the air duct and defines an interior fairing space in which at least a forward portion of the air duct is positioned. The fairing has a forward end which surrounds the air inlet of the air duct. At least one tube is provided having an inlet end in communication with a heated air chamber associated with an airframe structure of the aircraft, and an outlet end in communication with the interior fairing space. Heated air in the heated chamber is directed through the tube to the interior fairing space so as to provide icing protection to the air duct. The air scoop icing protection systems may advantageously be employed so as to provide icing protection to air scoops associated with an aircraft engine nacelle using the residual heat from the nacelle's inlet lip primary bleed air system.