• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    The invention relates to the heating of an aeronautical equipment intended to be disposed at the skin of the aircraft. The equipment (25) comprises a portion (30, 31) intended to be disposed at the level of the skin (27) of the aircraft and means for heating the part. According to the invention, the heating means comprise a thermodynamic loop comprising a closed circuit in which circulates a heat transfer fluid, the closed circuit comprising an evaporator (14) and an area in which condensation of the heat transfer fluid can take place in the appendix to warm it up. Outside the evaporator (14), the circuit in which the fluid flows is formed by a tubular channel with empty section. At least part of the equipment is made of a material with low thermal conductivity.
    公开号:FR3039509A1
    申请号:FR1501614
    申请日:2015-07-28
    公开日:2017-02-03
    发明作者:Romain Hodot；Claude Sarno；Frederic Martin；Jean Philippe Pineau
    申请人:Thales SA；
    IPC主号:
    专利说明:

    The invention relates to the reheating of aeronautical equipment intended to be disposed at the level of the skin of the aircraft.
    To ensure its mission, an aircraft comprises several equipment including flush portions or appendages protruding from the skin of the aircraft.
    These appendices or these flush portions belong, for example, to probes making it possible in particular to measure various aerodynamic parameters of the air flow surrounding the aircraft, such as, in particular, the total pressure, the static pressure, the temperature, or the incidence of the flow of air. air near the skin of the aircraft.
    The total pressure makes it possible, together with the static pressure, to determine the local velocity of the air flow in the vicinity of the probe. Other probes make it possible, for example, to measure the local incidence of an air flow.
    The incidence probes may include movable appendages for orientation in the axis of the airflow surrounding the probe. The orientation of the probe makes it possible to determine the incidence of the air flow. Other impact probes can be equipped with fixed appendages equipped with several pressure taps.
    The pressure difference measured between these pressure taps also makes it possible to determine the incidence of the air flow surrounding the probe. Other equipment such as cameras also need to be installed flush or prominent with respect to the skin of the aircraft, for example on nacelles known in the English literature as the "pod".
    When flying at high altitude, the aircraft may encounter icing conditions.
    Specifically frost can form on the skin and appendages of the aircraft. The appearance of frost is particularly problematic for aerodynamic probes whose profiles can be modified by frost and whose pressure ports can be obstructed.
    Measuring instruments mounted on nacelles can also be disturbed by the appearance of frost.
    One way to prevent frost formation is to heat the appendages.
    Currently reheating is in most cases achieved by means of electrical resistors embedded in the appendages.
    Reheating is done by joule effect. For example to warm a probe of total pressure, it is necessary to dissipate several hundred watts.
    More precisely, this type of probe is formed of a mat bearing a closed tube at one of its ends and called pitot tube.
    Heating of the probe is carried out by means of a heating resistor in the form of a heating wire wound in the body of the probe, that is to say both in the mast and in the Pitot tube.
    To produce the heating wire, an electrical conductor comprising an alloy of iron and nickel coated with a mineral insulator such as alumina or magnesia is commonly used. The insulation is itself coated with a nickel sheath or "INCONEL (registered trademark)" for brazing the wire on the body of the probe.
    A method for producing such a probe is for example described in patent application FR 2 833 347 filed in the name of the applicant.
    The realization of the heating wire and its assembly in the probe require a series of complex and expensive operations.
    Another embodiment for heating a pitot tube probe had been contemplated in US Pat. No. 4,275,603.
    This document describes the use of a heat pipe bringing thermal energy around the tube. The return of the heat transfer fluid in the liquid state is provided in a porous material.
    This allows the probe to be arranged in any possible orientation on the skin of the aircraft.
    In practice, this solution has no industrial advantage because of the difficulty of inserting a porous material into a probe.
    The method for producing such a probe is at least as complex as that using a heating wire. The invention aims to propose a new heated probe and more generally an aeronautical equipment flush or having a heated external appendix whose realization is much simpler than that described in the prior art. For this purpose, the subject of the invention is aeronautical equipment intended to equip an aircraft, the equipment comprising a part intended to be disposed at a skin of the aircraft and means for reheating the part, characterized in that the reheating means comprise a thermodynamic loop comprising a closed circuit in which a heat-transfer fluid circulates, the closed circuit comprising an evaporator and an area in which condensation of the coolant in the appendage can take place in order to heat it, and in that out of the evaporator, the circuit in which the fluid flows is formed by a tubular channel with empty section and in that at least this part of the equipment is made of a low thermal conductivity material.
    According to other characteristics of the equipment according to the invention taken alone or in combination: the channel is configured so that the fluid circulates thereby by capillarity; - It comprises a coolant circulation pump; the tubular channel forms a single thermodynamic loop outside the evaporator; the tubular channel forms several thermodynamic loops in which the heat transfer fluid circulates in parallel outside the evaporator; the part is configured to be flush with the skin of the aircraft; the part is an appendage configured to be arranged prominently with respect to the skin of the aircraft; - It comprises a base for fixing the equipment on the skin of the aircraft, the appendix is disposed on one side of the base and the evaporator is disposed on a second side of the base, opposite on the first side; the material has a thermal conductivity at 50 ° C. of less than 100 W / m · K; the material has a thermal conductivity at 50 ° C., preferably less than 60 W / m · K; it comprises at least two materials of different conductivities; it comprises a material having a conductivity gradient; - It is made of titanium; it comprises an aerodynamic measurement probe.
    According to another aspect the invention also relates to a method of producing an aeronautical equipment as described above, the equipment comprising a body in which the tubular channel with empty section is produced, the method being characterized in that the body is achieved by an additive manufacturing process.
    Finally, the invention also relates to a data file stored on storage means and loadable in the memory of a processing unit associated with an additive manufacturing machine capable of manufacturing an object by superposing layers of material, characterized in that it comprises data of three-dimensional representation of the equipment as described above, so as to enable, when loaded into the memory of, and processed by, said processing unit, the manufacture of said equipment by said manufacturing machine additive. The invention will be better understood and other advantages will appear on reading the detailed description of an embodiment given by way of example, a description illustrated by the accompanying drawing in which: FIG. 1a schematically represents a thermodynamic loop that can reheat aeronautical equipment; Figure 1b schematically shows several thermodynamic loops that can heat aircraft equipment; FIG. 2 represents an aerodynamic probe intended to measure the total pressure and equipping an aircraft; Figures 3a and 3b show a mat and a pitot tube forming external parts of the probe of Figure 1; Figure 4 shows an exploded view of different components of the probe; Figures 5a and 5b show an aerodynamic probe for measuring the static pressure and equipping an aircraft.
    For the sake of clarity, the same elements will bear the same references in the different figures.
    Figure 1a schematically shows a thermodynamic loop 11 in which circulates in a closed circuit, a coolant.
    In this loop, the fluid can be in two phases: liquid 12 and steam 13.
    The latent heat of transformation between these two phases is exploited to transport thermal energy between an evaporator 14 and a condenser 15.
    This type of thermodynamic loop is widely used to cool electronic components dissipating heat during their operation.
    In general, a heat input, schematized by arrows 16, at the level of the evaporator 14, is transported by the fluid in the vapor phase 13 to the condenser 15 where the energy supply is returned to the surrounding environment.
    This restitution is schematized by arrows 17.
    The closed circuit also comprises a reservoir 18 containing coolant in the liquid state. The reservoir 18 is placed near the evaporator 14. The reservoir 18 feeds the loop 11 via the evaporator 14.
    Thus, as soon as a sufficient energy supply is captured by the evaporator 14, the fluid in the liquid state contained in the evaporator vaporizes. The overpressure due to evaporation pushes the fluid in the vapor state 13 towards the condenser 15 where the fluid returns to its liquid state to return towards the evaporator 14.
    In the present application, the thermodynamic loop 11 is used to heat a part of an on-board aeronautical equipment. On board an aircraft, many pieces of equipment have appendages prominent with respect to the skin of the aircraft or flush portions.
    This equipment can be aerodynamic probes, antennas, sensors ...
    These appendages or flush portions require reheating to allow their operation. This heating is particularly important for aerodynamic probes that have ports used as pressure taps.
    Reheating prevents the formation of frost that could obstruct these orifices.
    The incidence probes, having a pallet intended to orientate themselves in the bed of the air flow surrounding the probe, are also sensitive to frost that could form on the pallet and change its shape, thus causing a bad measurement, even even a blockage of the pallet.
    FIG. 1b schematically represents two thermodynamic loops 11a and 11b in which the coolant circulates in parallel out of an evaporator 14 common to the various loops.
    These different loops 11a and 11b make it possible to heat more specifically different zones, forming condensers 15a and 15b, an appendage or part of an aeronautical equipment. The invention can of course be implemented for more than two thermodynamic loops.
    FIG. 2 represents an aeronautical probe 25 making it possible to measure the total pressure of an air flow surrounding the skin 27 of an aircraft.
    The probe 25 is intended to be fixed through an opening 26 made in the skin 27 of the aircraft.
    In FIG. 2, the skin 27, at its opening 26, is shown at a distance from the probe 25.
    The probe 25 comprises a Pitot tube 30 and a mast 31 carrying the Pitot tube 30.
    The pitot tube 30 and the mast 31 are external to the skin 27.
    The probe 25 also comprises an inner portion of the skin 27 having a pneumatic connector 32 for pneumatically connecting the pitot tube 30 to a pressure sensor located inside the fuselage of the aircraft.
    The probe 25 is positioned on the skin 27 of the aircraft so that the Pitot tube 30 is oriented substantially along a longitudinal axis of the aircraft, out of the boundary layer, so that the direction of flow, materialized by an arrow 33, substantially faces an inlet orifice 34 located at a first end 35 of the Pitot tube 30.
    A second end 36 of the pitot tube 30, opposite the end 35, is closed so as to create a stopping point in the stream of air taken from the flow and entering the tube 30 through its orifice 34.
    At the end 36 of the tube, a pneumatic channel, not shown in Figure 2, opens into the tube 30 to form a pressure point at which one seeks to measure the air pressure.
    The pneumatic channel is for example connected to a pressure sensor or other pressure measuring device such as for example a flow meter.
    The pressure sensor makes it possible effectively to measure the pressure of the air inside the tube 30 at its obstructed end 36.
    The pressure sensor may belong to the probe 25 or be remote. In this case, the pressure sensor is connected to the probe 25 by means of a pipe and the pneumatic connector 32. At the end 36, the tube 30 comprises one or more unrepresented bleed holes and allowing the evacuation of the water penetrating inside the tube 30.
    Apart from the purge hole or holes, the section of which is small relative to that of the tube 30, the tube 30 is closed at its end 36.
    The pressure measured at this end therefore represents the total pressure R of the air flow.
    The mast 31 carries the pitot tube 30 at its second end 36.
    The pitot tube 30 has a substantially cylindrical shape and the mast 31 an elongate shape. The mast 31 has for example a wing shape whose intrados and extrados may be symmetrical.
    The probe 25 may comprise other pressure taps, for example pressure taps placed on the mast 31 or around the tube 30 on its cylindrical part and making it possible to define the local incidence of the flow with respect to the probe Or measuring the static pressure of the flow.
    The probe 25 includes fixing means for attaching the probe 25 to the skin 27 of the aircraft.
    These means comprise for example a base 38 formed by a shoulder intended to come into contact with the skin 27.
    Screws arranged around the opening 26 immobilize the base 38 relative to the skin 27.
    In the example shown, the pitot tube 30 is fixed relative to the skin 27 of the aircraft.
    It is of course possible to mount the pitot tube 30 on a mobile mast such as a pallet that can be oriented in the axis of the flow, as described for example in the patent published under No. FR 2,665,539. and filed August 3, 1990. The base 38 then comprises a pivot connection allowing rotation of the mast 31 relative to the skin 27 about an axis perpendicular to the skin 27.
    Thus, as the local incidence of flow in the vicinity of the probe changes, the orientation of the pitot tube 30 follows this incidence in order to always cope with the flow.
    The measurement of total pressure Pt is improved during variations in local incidence of flow along the skin 27 of the aircraft. The evaporator 14 and the tank 18 are disposed inside the fuselage of the aircraft on one side of the base 38.
    The condenser 15 is formed of a channel disposed in the mast 31 and in the pitot tube 30.
    Heating means make it possible to supply thermal energy to the evaporator 14.
    These means comprise, for example, an electric heating resistor 40 arranged around the evaporator 14.
    Any other means allowing the supply of heat to the evaporator can also be implemented within the scope of the invention, such as, for example, the passage of a flow of hot air along the external walls of the evaporator 14. .
    Of course, other means can be envisaged as will be described in more detail later.
    It is possible to place in the appendix, a temperature sensor for measuring its temperature to control the heating means.
    Alternatively, a temperature measurement of the fluid in the evaporator 14 gives an image of the temperature of the appendix. The use of a thermodynamic loop for heating the probe 25 and more generally an aeronautical appendix has the advantage of facilitating the regulation of the temperature of the appendix by controlling the heating means delocalized inside the skin of the aircraft in the vicinity of the appendix.
    The fluids generally used as heat transfer fluids in a two-phase thermodynamic loop can have high latent heat of transformation, which makes it possible to reduce the flow of fluid in the loop for the same heat exchange.
    The reduction of the flow makes it possible to reduce the pressure losses in the loop. By way of example, it is possible to use methanol as heat transfer fluid.
    In what has just been described, the fluid flows in a tubular channel 39 with an empty section between the evaporator 14 and the condenser 15, in the condenser 15 itself, and between the condenser 14 and the evaporator 14.
    In other words, out of the evaporator 14, the circuit in which the fluid circulates is formed by the tubular channel 39 with empty section.
    Tubular channel with empty section means a channel comprising no filling except for the fluid of course.
    In particular, no porous material is present in the tubular channel 39. The internal walls of the tubular channel 39 are smooth to facilitate the circulation of the fluid and to limit the pressure losses.
    Figures 3a and 3b show an example of arrangement of the channel 39 fitted to the outer parts of the probe 25 and in which circulates the heat transfer fluid for heating these external parts.
    The mast 31 and the tube 30 both comprise an envelope, 41 for the mast and 42, for the tube 30.
    The pneumatic channel used for the measurement of pressure circulates in the envelope 41. The channel 39 is made in the respective envelopes.
    In the channel 39, the fluid circulating therein is likely to condense to heat the corresponding envelope or a part thereof as needed.
    More specifically, another advantage related to the realization of the tubular channel 39 empty section is the self-adaptability of heat exchange at the probe.
    Indeed, the coefficient of exchange between the fluid and the wall, coefficient of condensation, is related to the temperature gradients between the fluid and the wall.
    Thermal exchanges are more important in the colder zones of the probe 25. These colder zones correspond to the zones of the envelopes where external cooling is the most important.
    This makes it possible to obtain a better temperature homogeneity of the probe.
    Figure 3a shows the mast 31 and the pitot tube 30 in profile. One distinguishes an example of routing of the channel 39 in the corresponding envelopes.
    FIG. 3b shows in section the mast 31 in a plane parallel to the skin 27 in the vicinity of the opening 26.
    In its path, the channel 39 can be broken down into three parts 39a, 39b and 39c succeeding each other.
    After leaving the evaporator 14, the fluid circulates in the portion 39a made in the envelope 41. The portion 39a can meander in the envelope 41 between the leading edge 31a and the trailing edge 31b of the mast 31.
    Then the channel 39 winds in the envelope 42 by its part 39b. The path of the portion 39b is for example helical around the internal cavity of the pitot tube 30 at the bottom of which the total pressure is measured.
    The channel 39 continues its course in its part 39c by circulating again in the envelope 41 of the mast 31.
    As for part 39a, part 39c can meander in envelope 41 between leading edge 31a and trailing edge 31b.
    The definition of the path of the channel 39 is performed depending on the areas of the probe that should be warmed preferentially.
    In the example shown, the channel 39 winds in the appendix forming a single loop out of the evaporator 14.
    It is also possible to make in the appendix several loops in which the coolant flows in parallel out of the evaporator 14, as shown schematically in Figure 1b. The self-adaptation of the heat exchange to the actual temperature of the outer walls of the probe 25 allows a more tolerant path definition than for a probe heated directly by an electrical resistance.
    The section of the channel may be variable along its path in the mast 31 and in the Pitot tube 30.
    The circulation of the fluid in the channel 39 can be ensured by means of a circulation pump 45 disposed upstream of the evaporator 14. The circulation pump 45 is advantageously disposed inside the skin 27 of the aircraft.
    Alternatively, it is possible to dispense with this circulation pump 45 by configuring the section of the channel 39 in its various parts 39a to 39c so that the fluid circulates in its liquid phase by capillarity.
    Such a mode of circulation requires relatively small sections.
    In order to maintain a sufficient overall rate, the channel 39 may comprise zones placed in parallel.
    It is advantageous to make the probe 25 and more generally any aeronautical equipment embodying the invention by implementing an additive manufacturing process for manufacturing the mechanical part or parts in which the channel 39 travels.
    This process is also known as 3D printing.
    It is indeed possible to use for example this technology to manufacture at least part of this equipment using a low thermal conductivity material. The heat input is indeed closer to the areas concerned.
    For example, a material having a thermal conductivity at 50 ° C. of less than 100 W / m · K and preferably less than 60 W / m · K can be used.
    For example, titanium can be used.
    Of course other materials can be envisaged.
    Likewise, it is possible to envisage making this equipment using, for example, a material having a conductivity gradient or two or more materials having different thermal conductivities depending on the zones of this equipment.
    FIG. 4 is an exploded view of several mechanical parts which, once assembled, form the probe 25.
    A body 47 forms the base 38 and the envelopes 41 and 42. The channel 39 can be directly made in the body 47 by additive manufacturing.
    The body 47 can remain open at its trailing edge, for example to have in the body, the pneumatic channels for performing the total pressure measurement.
    Alternatively these channels can also be realized by the additive manufacturing process.
    The trailing edge 31a of the mast 31 and the end 36 of the pitot tube can be closed by means of a plug 48 which can be made by any type of manufacturing process.
    The shapes of the plug 48 are simpler than those of the body 47. It is for example possible to make the plug 48 by molding.
    Additive manufacturing is of course also usable for the plug 48.
    A support 49 can complete the probe 25.
    The support 49 may be used to carry the pneumatic connector in a first portion 49a and the evaporator 14 in a second portion 49b.
    The support is assembled to the body 47 by the base 38.
    Figures 5a and 5b show another aerodynamic probe 60.
    More specifically, the probe 60 forms aeronautical equipment comprising a portion 61 intended to be flush with the skin 27 of the aircraft.
    Figure 5a is a view in the plane of the skin 27 in the vicinity of the probe 60.
    Figure 5b is a sectional view perpendicular to the plane of the skin 27.
    Part 61 is for example in the form of a disc from closing an orifice 62 of the skin 27. The orifice 62 is provided to receive the portion 61 which is fixed by screw on the skin 27.
    The probe 60 is for example a static pressure probe having one or more pressure taps 63 formed by channels opening substantially perpendicular to the skin 27.
    The channel 39 circulates in the part 61. The channel winds around the pressure taps 63 in order to warm the part 61 and to prevent the pressure taps from being blocked by frost.
    In this embodiment, the channel 39 may also form a single loop or several parallel loops out of the evaporator 14.
    The probe 60 also includes an inner portion 65 to the skin 27. The inner portion 65 accommodates a pressure sensor connected to the pressure taps to measure the static pressure of the air flowing along the skin 27 The inner part 65 can also accommodate the evaporator 14 and the reservoir 18.
    As for the probe 25, the portion 61 is advantageously carried out by implementing an additive manufacturing process. The invention thus also relates to a data file stored on storage means and loadable in the memory of a processing unit associated with an additive manufacturing machine capable of manufacturing an object by superposing layers of material, which comprises data three-dimensional representation of the equipment as described above, so as to allow, when loaded in the memory of, and processed by, said processing unit, the manufacture of said equipment by said additive manufacturing machine. Other embodiments may still be contemplated.
    权利要求:
    Claims (14)
    [1" id="c-fr-0001]
    1. Aircraft equipment intended to equip an aircraft, the equipment (25) comprising a portion (60; 30; 31) intended to be disposed at a skin (27) of the aircraft and means for heating the aircraft. part, characterized in that the heating means comprise a thermodynamic loop comprising a closed circuit in which circulates a heat transfer fluid, the closed circuit comprising an evaporator (14) and a zone (15) in which condensation of the fluid can take place coolant in the appendix to heat it, and in that out of the evaporator (14), the circuit in which the fluid flows is formed by a tubular channel (39) with empty section and in that at least this part the equipment is made of a material with low thermal conductivity, having a thermal conductivity at 50 ° C of less than 100W / mK
    [2" id="c-fr-0002]
    2. Aeronautical equipment according to claim 1, characterized in that the channel (39) is configured so that the fluid circulates there by capillarity.
    [3" id="c-fr-0003]
    3. Aeronautical equipment according to claim 1, characterized in that it comprises a circulation pump (45) of the heat transfer fluid.
    [4" id="c-fr-0004]
    Aeronautical equipment according to one of the preceding claims, characterized in that the tubular channel (39) forms a single thermodynamic loop (11) out of the evaporator (14).
    [5" id="c-fr-0005]
    5. Aeronautical equipment according to one of claims 1 to 3, characterized in that the tubular channel (39) forms a plurality of thermodynamic loops (11a, 11b) in which the coolant flows in parallel from the evaporator (14).
    [6" id="c-fr-0006]
    6. Aircraft equipment according to one of the preceding claims, characterized in that the portion (61) is configured to be flush with the skin (27) of the aircraft.
    [7" id="c-fr-0007]
    7. Aeronautical equipment according to any one of claims 1 to 5, characterized in that the portion is an appendix (30, 31) configured to be arranged prominently relative to the skin (27) of the aircraft.
    [8" id="c-fr-0008]
    8. Aircraft equipment according to claim 7, characterized in that it comprises a base (38) for fixing the equipment on the skin (27) of the aircraft, in that the appendix (30, 31) is disposed on a first side of the base (38) and in that the evaporator (14) is disposed on a second side of the base (38), opposite the first side.
    [9" id="c-fr-0009]
    Aeronautical equipment according to one of the preceding claims, characterized in that the material has a thermal conductivity at 50 ° C of less than 60W / m.K.
    [10" id="c-fr-0010]
    10. Aircraft equipment according to one of claims 1 to 8, characterized in that it comprises at least two materials of different conductivities.
    [11" id="c-fr-0011]
    11. Aircraft equipment according to one of claims 1 to 8, characterized in that it comprises a material having a conductivity gradient.
    [12" id="c-fr-0012]
    12. Aeronautical equipment according to any one of the preceding claims, characterized in that it is made of titanium.
    [13" id="c-fr-0013]
    Aircraft equipment according to any one of the preceding claims, characterized in that it comprises an aerodynamic measuring probe (30).
    [14" id="c-fr-0014]
    14. A method of producing aeronautical equipment according to one of the preceding claims, the equipment comprising a body (47) in which is formed the tubular channel (39) with empty section, the method being characterized in that the body (47) is achieved by an additive manufacturing process.
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    US20160304210A1|2014-10-15|2016-10-20|Rosemount Aerospace Inc.|One-piece air data probe|
    ITUB20151085A1|2015-05-28|2016-11-28|Alenia Aermacchi Spa|Nacelle for aircraft engine with de-icing system using a two-phase fluid.|
    FR3039509B1|2015-07-28|2017-12-22|Thales Sa|HEATING FOR AERONAUTICAL EQUIPMENT FOR AN AIRCRAFT|FR3039509B1|2015-07-28|2017-12-22|Thales Sa|HEATING FOR AERONAUTICAL EQUIPMENT FOR AN AIRCRAFT|
    FR3039508B1|2015-07-28|2017-09-01|Thales Sa|HEATING FOR AIRCRAFT EQUIPMENT|
    FR3094345B1|2019-03-27|2021-03-05|Thales Sa|AERONAUTICAL EQUIPMENT FOR AN AIRCRAFT|
    法律状态:
    2016-07-29| PLFP| Fee payment|Year of fee payment: 2 |
    2017-02-03| PLSC| Publication of the preliminary search report|Effective date: 20170203 |
    2017-07-31| PLFP| Fee payment|Year of fee payment: 3 |
    2018-07-27| PLFP| Fee payment|Year of fee payment: 4 |
    2019-07-31| PLFP| Fee payment|Year of fee payment: 5 |
    2020-07-31| PLFP| Fee payment|Year of fee payment: 6 |
    2021-07-29| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1501614A|FR3039509B1|2015-07-28|2015-07-28|HEATING FOR AERONAUTICAL EQUIPMENT FOR AN AIRCRAFT|FR1501614A| FR3039509B1|2015-07-28|2015-07-28|HEATING FOR AERONAUTICAL EQUIPMENT FOR AN AIRCRAFT|
    GB1612976.9A| GB2541995B|2015-07-28|2016-07-27|Aeronautic Equipment|
    US15/220,417| US10329024B2|2015-07-28|2016-07-27|Non conducting material|
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