69results about "De-icing equipments" patented technology

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 present invention relates to an apparatus and method for detecting the presence of water or ice on a structure, for example on the surface of an aircraft. A plurality of heaters (202-214) are thermally coupled to a structure (for example on the back of a wing) (104) in order to detect the presence of water or ice on the structure. The heaters are arranged adjacent one another from a GC region of a leading edge (106) of the structure (that is exposable to an impinging airflow) and extending aft of the leading edge of the structure. The heaters, which may be controlled individually, are supplied power that is sufficient to heat the surface of the structure to substantially the same temperature. A controller senses the power required for the heaters to achieve the same surface temperature at the respective regions. By comparing the power consumed by a heater that is aft of the fore-most heater (214), and the power consumed by a heater fore of the aft-most heater (210), a determination of the presence of water or ice can be made if the power consumed by the heater that is aft of the fore-most heater is different to the power consumed by the heater that is fore of the aft-most heater.

The present application provides a wing hot air anti-icing control system and an air anti-icing pressure control method. The system comprises a wing anti-icing valve, a wing anti-icing pressure sensor, a wing anti-icing temperature sensor, a wing anti-icing ground temperature monitoring sensor, a wing anti-icing monitoring pressure sensor, a wing anti-icing controller, a bleed air fan valve and a flute type pipe. The wing anti-icing controller comprises independent control channels and monitoring channels, and the wing anti-icing valve comprises independent valve shutdown circuits and valve adjustment circuits. The present application can realize ground and air anti-icing, and the anti-icing is more accurate according to the wing anti-icing pressure sensor than the anti-icing according to the reference upstream bleed air temperature. In the case of single-engine bleed air anti-icing, bleed air failure or performance decline, etc., the anti-icing or maximum anti-icing can still be performed under the condition of ensuring the optimal aircraft. Therefore, the present application can not only effectively reduce energy consumption, improve economy and safety, but also has good anti-icing effect and good universality.

This invention discloses an intelligent, controllable, and self-healing honeycomb torsion rotor and its control method, belonging to the field of rotorcraft technology, applicable to helicopters, multi-rotor UAVs, electric vertical takeoff and landing aircraft, and any aircraft whose primary power source is a rotor. The rotor of this invention includes a rotor hub and at least two blades. The blades have a multi-layered composite structure, including: an aerodynamic skin, an intelligent honeycomb sandwich structure (composed of piezoelectric composite honeycomb cells and MFC actuators, embedded with bistable cells), an internal vortex tube heat exchange channel, a self-healing functional layer (microcapsules and / or dynamic covalent bond system), a surface-programmable electronic paper unit, and a control system. This invention utilizes the rotor's centrifugal force for passive air intake, and vortex tube heat exchange to achieve cold and hot separation; the hot air can be used for de-icing and forming a Coanda effect air cushion; active blade torsion and zero-power state maintenance are achieved through piezoelectric intelligent materials and a bistable structure; and the self-healing functional unit enables automatic healing of microcracks. This invention integrates active deformation, thermal management, structural self-repair, and air cushion lift enhancement, solving the shortcomings of existing helicopter rotor structures, such as complex structure, limited efficiency, difficult thermal management, and lack of self-repair capability.

A data communication method for a de-icing system for a turbine engine is disclosed. The turbine engine includes a fixed reference frame and a rotating reference frame. The system includes a non-contact rotating power transformer (RTU), the primary side of which is constrained by the fixed reference frame, and the secondary side of which is constrained by the rotating reference frame. The fixed reference frame includes an ice protection and control unit (IPC) and a power supply device (DC / AC) for supplying AC power to the primary side. The IPC transmits commands to the DC / AC to adjust the level of power supplied to the primary side. The rotating reference frame includes a propeller (1) and a switch, with a plurality of resistors fixed to the propeller. The switch is used to sequentially distribute the power supplied to the secondary side to the resistors. The protection and control unit (IPC) is programmed to transmit switching commands to the switch via the power supply unit (DC / AC) and the transformer (RTU) by issuing instructions (Ic1, Ic2, Ic3, Ic4, Ic5, Intl) to the power supply unit (DC / AC) to encode the voltage wave sequence (To) applied to the primary side. The switch is programmed to decode the voltage wave sequence (To) recovered on the secondary side.

Ice may form along the leading edge of an aircraft wing. A pneumatic deicing system may be configured to crack and dislodge ice along the leading edge of the wing. The pneumatic deicing system may comprise a deicing boot assembly having a deicing boot attached to the leading edge, and a gas generator fluidly coupled to the deicing boot assembly. The gas generator may comprise a propellant and may decompose the propellant, liberating a compressed gas. The compressed gas may be directed to the deicing boot assembly, inflating the deicing boot, which may crack and dislodge the ice.

The application discloses a double FBG ice accumulation and deicing detection system embedded between an airplane primer and a topcoat and a signal processing method thereof, and belongs to the technical field of airplane state monitoring. The method comprises the following steps: embedding a double FBG sensor structure containing a temperature compensation grating and a strain sensing grating in an airplane wing coating and performing temperature-strain decoupling to obtain a pure strain component; performing intelligent filtering processing on the component to obtain a smooth strain signal; based on the signal and a change rate thereof, in combination with a preset threshold, dynamically identifying ice layer growth, peeling, stability or residual state; and finally, quantitatively evaluating a deicing effect based on sequence similarity analysis. A corresponding system comprises a sensing module, a signal demodulation module and a data processing and decision module. Through in-coating embedded sensing and multi-stage signal processing, the application realizes high-precision and dynamic monitoring of the airplane ice accumulation and deicing process, and effectively solves the problems of temperature interference and noise influence.

Aircraft propeller blade de-icing system comprising a primary measuring winding. Aircraft propeller blade (12) de-icing system (10) comprising an H-bridge type inverter device (40); a rotating transformer (22) comprising a main primary winding (26), a primary measuring 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 a phase-shifted manner and using the value of the voltage (Vpm) across the terminals of the primary measuring winding measured during a measurement interval of a predetermined duration and which begins at the instant (t11,t33) when the inverter device transitions from a controlled state to a short-circuit state. Figure for the abridged version: Fig. 1.

A heated leading edge structure (14) for an aircraft (10) including a leading edge panel (20) having an outer surface (32) configured to contact 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 arrangement (24) configured to heat the leading edge panel (20); wherein the heating arrangement (24) includes an indirect heating device (44) mounted on the reinforcing member (22) and configured to heat the leading edge panel (20) by conduction through the reinforcing member (22).

The present application relates to the technical field of aircraft icing prediction, and particularly relates to a hierarchical prediction method and system for aircraft icing, comprising the steps of: obtaining hierarchical indexes in a flight process; performing hierarchical processing on the hierarchical indexes to obtain at least one hierarchical membership degree; obtaining cloud top temperature in the flight process; determining a cloud top temperature membership degree according to the cloud top temperature and a corresponding cloud top temperature membership degree function; forming at least two groups of membership degrees according to the at least two hierarchical membership degrees and the cloud top temperature membership degree; calculating at least two initial icing potentials according to the at least two groups of membership degrees and an initial icing potential model; calculating at least two final icing potentials according to the at least two initial icing potentials; and outputting corresponding hierarchical prediction results according to the at least two final icing potentials. The present application can greatly improve the accuracy and efficiency of aircraft icing prediction with the aid of fuzzy logic.

The application discloses a power design method and device of a wing electrothermal deicing system and a storage medium, and relates to the technical field of electrothermal deicing. The application measures the ice sample shedding time of a plurality of target positions on the surface of a target airfoil through icing wind tunnel test. The spatial distribution difference of the ice adhesion strength is accurately characterized through the ice shedding time spatial distribution curve, and then a differentiated power calculation model is used to configure an independent power calculation formula for each target position. After the engineering constraint partition, the heating is started synchronously, the ice adhesion strength difference is accurately compensated by the partition power density difference, the deicing asynchronization problem caused by the traditional uniform power or experience is solved, the synchronous deicing of the plurality of positions of the wing is realized, the stable flight aerodynamic performance is ensured, and the invalid energy consumption can be reduced and the energy utilization rate can be improved.

A gas turbine engine inlet anti-ice system includes a nacelle having an inlet lip, an anti-ice air supply duct, and a flow deflector. The inlet lip has an annular anti-icing chamber formed therein that is defined by an inner surface of the inlet lip and an aft wall that has an aft wall opening formed therein. The anti-ice air supply duct is configured to direct a flow of pressurized air into the aft wall opening. The flow deflector is coupled to the aft wall and extends over the aft wall opening. The flow deflector has an inner surface, an outer surface, a closed end, and an open end. The flow deflector and aft wall define a flow channel that extends between the closed end and the open end and in a direction that is tangential to the aft wall opening.

An assembly (54) for an aircraft propulsion system (20) includes a nacelle inlet structure (34) which extends axially along and circumferentially around a centerline (28). The nacelle inlet structure (34) includes an inlet lip (62), an inner barrel (58), an outer barrel (60) and a plurality of structure segments (160). The inlet lip (62) forms a leading edge (86) of the nacelle inlet structure (34). The inner barrel (58) projects axially aft away from the inlet lip (62). The outer barrel (60) projects axially aft away from the inlet lip (62). The outer barrel (60) is radially outboard of and axially overlaps the inner barrel (58). Each of the structure segments (160) includes an exterior skin (162), a mount (164) and an electric heater (166) configured to heat the exterior skin (162). The exterior skin (162) forms a respective circumferential section of the inlet lip (62). The mount (164) is connected to the exterior skin (162). A fastener (190) extends radially through the mount (164) and attaches a respective one of the structure segments (160) to the inner barrel (58).

Aeronautical measuring instrument and associated manufacturing method. The present invention relates to an aeronautical measuring instrument (10) comprising a body (22) and a heating element (40), the body (22) defining a region of interest (25) intended to be exposed to aerodynamic flows. The instrument (10) is characterized in that the heating element (40) is encapsulated in the body (22) and configured to heat the region of interest (25). Figure for the abstract: Figure 1

An assembly for aircraft to prevent bird strikes by deterring the birds using an 18 kHz signal. The assembly can be installed on the front of the aircraft. Each assembly includes a radio transmitter configured to generate an 18 kHz signal with a 5-10-mile range.

The present application relates to the technical field of aviation deicing engineering, and particularly relates to a bionic photothermal-phase change composite interface and a preparation method and application thereof. The bionic photothermal-phase change composite interface is composed of a hemispherical cap microcolumn array and an inclined groove coupling substrate, and a bonding phase-photothermal component-n-nineteen alkane composite coating. The preparation is performed by 3D printing to prepare the substrate, pretreatment, preparation of a suspension, multiple thin layer spraying and solidification to obtain the finished product. The interface is heated to 33 DEG C within 120s under light to trigger phase change, realize super-hydrophobic-super-hydrophilic reversible switching, delay icing in the super-hydrophobic state, low ice adhesion, directional drainage through triple gradient in the super-hydrophilic state, stable performance through multiple durability tests, no residual water in the wing simulation, and providing a new scheme for aviation low-energy consumption deicing, and having engineering and migration.

The present invention relates to a method and arrangement for detecting presence of possible icing conditions of an unmanned aerial vehicle. The method obtains (206) ambient temperature information of the UAV, records (201) at least one diffraction image (25a) comprising information of cloud droplets (25) situated in an open measurement space. The method further transfers (202) the recorded diffraction images (25a) as respective holograms to a computer (3). The method further applies (205) a trained neural network for recognizing the cloud droplets (25), where an input of the trained neural network is the at least one hologram and where an output of the trained neural network comprises particle size information of the cloud droplets (25), and calculating (208) a presence of possible icing conditions based on the obtained ambient temperature information of the UAV and the particle size information of the cloud droplets (25).

The invention discloses an aircraft deicing device, and the device comprises a vibration module which is used for converting the electric energy provided by an external power supply into mechanical vibration through an inverse piezoelectric effect; the vibration transmission module comprises a transmission end and an action end, the transmission end is in coupling connection with the driving end of the vibration module, and the action end is tightly attached to the surface of the aircraft structure to be deiced and used for transmitting mechanical vibration generated by the vibration module and acting on the surface of the aircraft structure to be deiced. The ice on the surface of the aircraft structure is removed. Efficient, low-energy-consumption and rapid deicing on the surface of the aircraft can be achieved.

The present disclosure provides for pneumatic deicer boot assemblies (10), and related methods of fabrication and use. More particularly, the present disclosure provides for bartack stitched pneumatic deicer boot assemblies for aircraft or the like, with the bartack stitched pneumatic deicer boot assemblies having an inflatable carcass formed by two layers (e.g., a stretchable layer (12) and a non-stretchable layer (14)) stitched together using bartack stitches (16).

An aircraft wing assembly with a fixed wing and a wing tip device are disclosed. The wing tip device is moveable between a flight configuration and a ground configuration. In the ground configuration the span of the aircraft wing assembly is reduced. The fixed wing includes a fixed wing lug and the wing tip device comprises a wing tip lug. A locking pin is provided. In the flight configuration the fixed wing lug and wing tip device lug are aligned such that the locking pin may be inserted through the fixed wing lug and wing tip device lug into a locking position to lock the wing tip device in the flight configuration. The locking pin is hollow and includes a heating element.

This fan rotor comprises a hub, a plurality of blades (56), a system (62) for de-icing the blades (56), and a pitch change mechanism (70) for changing the pitch of the blades (56). The pitch change mechanism (70) comprises a control actuator (74) and a pump (212), actuated by an electric motor (214), for circulating a fluid within a circuit (120) supplying the control actuator (74). The fan rotor further comprises a rotary transformer (246) for transmitting an electric current by electromagnetic induction between a primary electric circuit (242) in relation to which the fan rotor (54) is rotatably mounted and a secondary electric circuit (244) which is integral with the hub and is configured to supply the electric motor (214) and the de-icing system (62) with said electric current.

The invention relates to a system (101) for discriminating the size of supercooled droplets for an aircraft. The system (101) comprises at least one laser source (103) configured to emit a laser beam (105) and a focusing device (107) for focusing the laser beam (105) into a cloud (109) of supercooled droplets. The laser source (103) is configured to generate an optical back-injection interference signal from a reflection of the laser beam (105) into the cloud (109) of supercooled droplets, and a processing unit (113) is configured to acquire the interference signal and analyze the amplitude variations of said interference signal using a convolutional neural network so as to classify the droplets in the cloud (109) according to their size. Figure for the abstract: Fig. 3

This invention discloses a system and method for de-icing aircraft propellers, relating to the field of de-icing device technology. The system includes a de-icing control module and a de-icing circuit breaker. The remote circuit breaker receives a standby command, sends it to the de-icing control module, and allows the de-icing control module to power on. The de-icing control module receives the standby command from the remote circuit breaker, powers on, and can distribute energy to a portion of the multiple de-icing units in turn. The method includes de-icing control steps, which include the remote circuit breaker receiving a standby command, sending it to the de-icing control module, and allowing the de-icing control module to power on. The de-icing control module receives the standby command from the remote circuit breaker, powers on, and distributes energy to a portion of the multiple de-icing units in turn. By distributing energy to a portion of the multiple de-icing units in turn through the de-icing control module and the de-icing circuit breaker, de-icing control is achieved.

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