Patent Publication Number: US-9849993-B2

Title: First equipment heating device

Description:
The invention relates to an aircraft, and more particularly to the heating of a first piece of aeronautic equipment intended to be arranged at the skin of the aircraft. 
     To perform its mission, an aircraft comprises several pieces of equipment comprising parts flush with or appendages protruding from the skin of the aircraft. 
     These appendages or these flush parts for example belong to probes in particular making it possible to measure different aerodynamic parameters of the airflow surrounding the aircraft, in particular the total pressure, static pressure, temperature, or incidence of the airflow near the skin of the aircraft. 
     The total pressure, combined with the static pressure, makes it possible to determine the local speed of the airflow near the probe. 
     Other probes for example make it possible to measure the local incidence of an airflow. 
     The incidence probes may comprise moving appendages intended to be oriented in the axis of the airflow surrounding the probe. 
     The orientation of the probe makes it possible to determine the incidence of the airflow. 
     Other incidence probes may be equipped with stationary appendages equipped with several pressure taps. 
     The pressure difference measured between these pressure taps makes it possible to determine the incidence of the airflow surrounding the probe. 
     Other pieces of equipment such as cameras also must be installed flush or protruding relative to the skin of the aircraft, for example on pods. 
     During flight at high altitudes, the aircraft may encounter freezing conditions. 
     More specifically, ice may form on the skin and appendages of the aircraft. The appearance of ice is particularly problematic for the aerodynamic probes, the profiles of which may be modified by ice and the pressure tap orifices of which may be obstructed. 
     The measuring instruments mounted on pods may also be disrupted by the appearance of ice. 
     One solution making it possible to avoid ice formation consists of heating the appendages. 
     Currently, in most cases, heating is done using electrical resistances embedded in the appendages. 
     The heating is done by Joule effect. For example, to heat a total pressure probe, it is necessary to dissipate several hundred watts. 
     More specifically, this type of probe is formed by a mast bearing a tube closed at one of its ends and called Pitot tube. 
     The heating of the probe is done using a heating resistance made in the form of a heating wire wound around the body of the probe, i.e., both in the mast and the Pitot tube. 
     To produce the heating wire, an electrical conductor is commonly used including an alloy of iron and nickel coated with an insulating material such as alumina or manganese. The insulator itself is coated with a nickel or Inconel sheath allowing brazing of 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 applicant&#39;s name. 
     The production of the heating wire and its assembly in the probe require a series of complex and costly operations. 
     Another embodiment to heat a Pitot tube probe had been considered in U.S. Pat. No. 4,275,603. 
     This document describes the use of a heat pipe contributing heat energy around the tube. The return of the heat transfer fluid to liquid state is ensured 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 due to the difficulty of inserting a porous material in a probe. 
     The method for producing such a probe is at least as complex as that implementing a heating wire. 
     The invention seeks to propose a new heated probe, and more generally a piece of aeronautic equipment that is flush or that has a heated outer appendage, the production of which is much simpler than that described in the prior art. 
     To that end, the invention relates to an aircraft provided with at least two pieces of aeronautic equipment, a first piece of equipment comprising at least one part intended to be arranged at a skin of the aircraft and means for heating the part, characterized in that the heating means comprise a thermodynamic loop comprising a closed circuit in which a heat transfer fluid circulates, the closed circuit comprising an evaporator associated with functional means of the second piece of equipment of the aircraft forming a heat source giving off heat during their operation and a zone in which a condensation of the heat transfer fluid can occur in the appendage to heat it, and in that outside the evaporator, the circuit in which the fluid circulates is formed by a tubular channel with an empty section. 
     According to other features of the aircraft according to the invention, considered alone or in combination:
         the functional means of the second piece of equipment of the aircraft forming a heat source are formed by a part or an electronic board thereof;   the functional means of the second piece of equipment of the aircraft forming a heat source are formed by an actuator thereof;   the functional means of the second piece of equipment of the aircraft forming a heat source are formed by a part or a piece of electrical equipment thereof;   the channel is configured for the fluid to circulate therein by capillarity;   it comprises a circulation pump for the heat transfer fluid;   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 protruding relative to the skin of the aircraft;   the first piece of equipment comprises a base intended to fasten the piece of equipment on the skin of the aircraft, the appendage is arranged on a first side of the base and the evaporator is arranged on a second side of the base, opposite the first side;   the first piece of equipment comprises an aerodynamic measuring probe.       

     According to another aspect, the invention also relates to a method for producing a first piece of aeronautic equipment of an aircraft as previously described, the equipment comprising a body in which the tubular channel with an empty section is produced, the method being characterized in that the body is produced using an additive manufacturing method. 
     Lastly, the invention also relates to a data file stored on storage means and able to be loaded in the memory of a processing unit associated with an additive manufacturing machine able to manufacture an object by superimposing layers of material, characterized in that it comprises data for three-dimensional depiction of the first piece of equipment for an aircraft as previously described, so as to allow, when it is loaded into memory of, and processed by, said processing unit, the manufacture of said piece of equipment by said additive manufacturing machine. 
    
    
     
       The invention will be better understood, and other advantages thereof will appear, upon reading the detailed description of one embodiment given as an example, this description being illustrated by the attached drawing, in which: 
         FIG. 1 a    diagrammatically shows a thermodynamic loop able to heat a piece of aeronautic equipment; 
         FIG. 1 b    diagrammatically shows several thermodynamic loops able to heat a piece of aeronautic equipment; 
         FIG. 2  shows an aerodynamic probe intended to measure the total pressure and equipping an aircraft; 
         FIGS. 3 a  and 3 b    show a mast and a Pitot tube forming outer parts of the probe of  FIG. 1 ; 
         FIG. 4  shows an exploded view of different component parts of the probe; 
         FIGS. 5 a  and 5 b    show an aerodynamic probe intended to measure the static pressure and equipping an aircraft; 
         FIG. 6  shows an example embodiment of an aerodynamic probe included in the makeup of an aircraft according to the invention. 
     
    
    
     For clarity reasons, the same elements bear the same references in the different figures. 
       FIG. 1 a    diagrammatically shows a thermodynamic loop  11  in which a heat transfer fluid circulates in a closed circuit. 
     In this loop, the fluid may assume two phases: liquid  12  and vapor  13 . 
     The latent transformation heat between these two phases is used to transport heat 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 contribution, diagrammed by arrows  16 , at the evaporator  14 , is transported by the fluid in vapor phase  13  toward the condenser  15 , where the energy contribution is returned to the surrounding environment. 
     This return is diagrammed by arrows  17 . 
     The closed circuit also comprises a reservoir  18  containing heat transfer fluid in liquid state. The reservoir  18  is arranged near the evaporator  14 . The reservoir  18  supplies the loop  11  via the evaporator  14 . 
     Thus, once a sufficient energy contribution is captured by the evaporator  14 , the fluid in liquid state contained in the evaporator vaporizes. The overpressure due to the evaporation pushes the fluid in vapor state  13  toward the condenser  15 , where the fluid regains its liquid state to return toward the evaporator  14 . 
     In the present application, the thermodynamic loop  11  is used to heat part of a piece of onboard aeronautic equipment. 
     Onboard an aircraft, many pieces of equipment have protruding appendages relative to the skin of the aircraft or flush parts. 
     These pieces of equipment may be aerodynamic probes, antennas, sensors, etc. 
     These appendages or these flush parts require heating to allow them to operate. This heating is particularly important for the aerodynamic probes, which have orifices used as pressure taps. 
     The heating makes it possible to avoid the formation of ice, which could obstruct these orifices. 
     The fire probes, which have a vane intended to be oriented in the bed of the airflow surrounding the probe, are also sensitive to the ice that may form on the vane and alter its shape, thus causing an incorrect measurement, or even blocking of the vane. 
       FIG. 1 b    diagrammatically shows two thermodynamic loops  11   a  and  11   b  in which the heat transfer fluid circulates in parallel outside an evaporator  14  shared by the different loops. 
     These different loops  11   a  and  11   b  more specifically make it possible to heat different zones, forming condensers  15   a  and  15   b,  of an appendage or part of a piece of aeronautic equipment. 
     The invention may of course be implemented for more than two thermodynamic loops. 
       FIG. 2  shows an aeronautic probe  25  making it possible to measure the total pressure of an airflow surrounding the skin  27  of an aircraft. 
     The probe  25  is intended to be fixed crossing through an opening  26  formed in the skin  27  of the aircraft. 
     In  FIG. 1 , 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  supporting the Pitot tube  30 . 
     The Pitot tube  30  and the mast  31  are outside the skin  27 . 
     The probe  25  also comprises a part inside the skin  27  including a pneumatic connector  32  allowing the pneumatic connection of the Pitot tube  30  to a pressure sensor situated inside the fuselage of the aircraft. 
     The probe  25  is positioned on the skin  27  of the aircraft such that the Pitot tube  30  is oriented substantially along a longitudinal axis of the aircraft, outside the boundary layer, so that the direction of the flow, embodied by an arrow  33 , is substantially across from an inlet orifice  34  situated at 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 stop point in the air taken from the flow and penetrating the tube  30  through its orifice  34 . 
     At the end  36  of the tube, a pneumatic channel, not shown in  FIG. 1 , opens in the tube  30  to form a pressure tap therein at which one seeks to measure the air pressure. 
     The pneumatic channel is for example connected to a pressure sensor or another pressure measuring device, for example a flowmeter. 
     The pressure sensor allows an effective measurement of the air pressure prevailing inside the tube  30  at its obstructed end  36 . 
     The pressure sensor can belong to the probe  25  or be offset. In this case, the pressure sensor is connected to the probe  25  using a hose and the pneumatic connector  32 . 
     At the end  36 , the tube  30  includes one or several drain holes, not shown, allowing the discharge of the water penetrating the inside of the tube  30 . 
     Aside from the bleed hole(s), which have a small section 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 Pt of the flow of air. 
     The mast  31  bears the Pitot tube  30  at its second end  36 . 
     The Pitot tube  30  has a substantially cylindrical shape and the mast  31  has an elongated shape. The mast  31  is for example in the shape of a wing, the concave and convex sides of which may be symmetrical. 
     The probe  25  may comprise other pressure taps, for example pressure taps arranged 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 relative to the probe  25  or measuring the static pressure of the flow. 
     The probe  25  comprises fastening means intended to fasten the probe  25  to the skin  27  of the aircraft. 
     These means for example comprise 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 illustrated example, the Pitot tube  30  is stationary relative to the skin  27  of the aircraft. 
     It is of course possible to mount the Pitot tube  30  on a moving mast, for example a vane that may be oriented in the axis of the flow, as for example described in the patent published under no. FR 2,665,539 and filed on Aug. 3, 1990. 
     The base  38  then comprises a pivot link allowing the rotation of the mast  31  relative to the skin  27  around an axis perpendicular to the skin  27 . 
     Thus, when the local incidence of the flow, near the probe  25 , evolves, the orientation of the Pitot tube  30  follows this incidence so as always to face the flow. 
     The total pressure measurement Pt is thereby improved during local incidence variations of the flow along the skin  27  of the aircraft. 
     The evaporator  14  and the reservoir  18  are arranged inside the fuselage of the aircraft on one side of the base  38 . 
     The condenser  15  is formed by a channel arranged in the mast  31  and in the Pitot tube  30 . 
     Heating means make it possible to contribute heat energy to the evaporator  14 . 
     These means for example comprise a heating electrical resistance  40  arranged around the evaporator  14 . 
     Any other means making it possible to contribute heat to the evaporator may also be implemented in the context of the invention, for example the passage of a hot air flow along the outer walls of the evaporator  14 . 
     Of course, other means may be considered, as will be described in more detail below. 
     It is possible to place a temperature sensor in the appendage, making it possible to measure its temperature to enslave the heating means. 
     Alternatively, a temperature measurement of the fluid in the evaporator  14  provides an image of the temperature of the appendage. 
     Using a thermodynamic loop to heat the probe  25 , and more generally an aeronautic appendage, has the advantage of facilitating the regulation of the temperature of the appendage by controlling the heating means delocalized inside the skin of the aircraft near the appendage. 
     The fluids generally used as heat transfer fluids in a diphasic thermodynamic loop can have high latent transformation heats, which makes it possible to reduce the fluid flow rate in the loop for a same heat exchange. 
     The reduction in flow rate makes it possible to reduce the pressure losses in the loop. 
     As an example, methanol may be used as heat transfer fluid. 
     In the situation described above, the fluid circulates 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, outside the evaporator  14 , the circuit in which the fluid circulates is formed by the tubular channel  39  with an empty section. 
     A tubular channel with an empty section refers to a channel not including any filling, other than the fluid, of course. 
     In particular, no porous material is present in the tubular channel  39 . The inner walls of the tubular channel  39  are smooth to facilitate the circulation of the fluid and limit the pressure losses. 
       FIGS. 3 a  and 3 b    show an example arrangement of the channel  39  equipping the outer parts of the probe  25  and in which the heat transfer fluid circulates that makes it possible to heat these outer parts. 
     The mast  31  and the tube  30  both comprise an enclosure,  41  for the mast, and  42  for the tube  30 . 
     The pneumatic channel used for the pressure measurement circulates in the enclosure  41 . The channel  39  is made in the respective enclosures. 
     In the channel  39 , the fluid that circulates therein is able to condense to heat the corresponding enclosure or part of that enclosure as needed. 
     More specifically, another advantage related to the production of the tubular channel  39  with an empty section is the auto-adaptation capacity of the heat exchanges at the probe. 
     Indeed, the exchange coefficient between the fluid and the wall, condensation coefficient, is related to the temperature gradients between the fluid and the wall. 
     The heat exchanges are greater in the coldest zones of the probe  25 . These coldest zones correspond to the zones of the enclosures where the outer cooling is greatest. 
     This makes it possible to obtain better homogeneity of the probe in terms of temperature. 
       FIG. 3 a    shows the mast  31  and the pitot tube  30  in profile. One example path of the channel  39  in the corresponding enclosures can be seen in this figure. 
       FIG. 3 b    shows the mast  31  in sectional view in a plane parallel to the skin  27  near the opening  26 . 
     Along its path, the channel  39  can be broken down into three parts  39   a ,  39   b  and  39   c , following one another. 
     After it leaves the evaporator  14 , the fluid circulates in the part  39   a  made in the enclosure  41 . The part  39   a  can wind in the enclosure  41  between the leading edge  31   a  and the trailing edge  31   b  of the mast  31 . 
     Next, the part  39   b  of the channel  39  winds in the enclosure  42 . The path of the part  39   b  is for example helical around the inner cavity of the Pitot tube  30  at the bottom of which the total pressure is measured. 
     The channel  39  continues its path in the part  39   c  while again circulating in the enclosure  41  of the mast  31 . 
     As for the part  39   a,  the part  39   c  can wind in the enclosure  41  between the leading edge  31   a  and the trailing edge  31   b  of the mast  31 . 
     The definition of the path of the channel  39  is done based on the zones of the probe that should preferably be heated. 
     In the illustrated example, the channel  39  winds in the appendage while forming a single loop outside the evaporator  14 . 
     It is also possible to produce several thermodynamic loops in the appendage, in which loops the heat transfer fluid circulates in parallel outside the evaporator  14 , as shown diagrammatically in  FIG. 1   b.    
     The auto-adaptation of the heat exchange to the actual temperature of the outer walls of the probe  25  allows a more tolerant definition of the path than for a probe heated directly by an electrical resistance. 
     The section of the channel may vary along its entire path in the mast  31  and in the Pitot tube  30 . 
     The circulation of the fluid in the channel  39  may be ensured using a circulation pump  45  arranged upstream from the evaporator  14 . The circulation pump  45  is advantageously arranged inside the skin  27  of the aircraft. 
     Alternatively, it is possible to do away with this circulation pump  45  by configuring the section of the different parts  39   a  to  39   c  of the channel  39  so that the fluid circulates in its liquid phase by capillarity. 
     Such a circulation mode requires relatively small sections. 
     In order to retain a sufficient overall flow rate, the channel  39  may comprise zones placed in parallel. 
     It is advantageous to produce the probe  25 , and more generally any piece of aeronautic equipment implementing the invention, by carrying out an additive manufacturing method to manufacture the mechanical part(s) in which the channel  39  travels. 
     This method is also known as 3D printing. 
     At this time, it is known to produce metal parts using this method. It is for example possible to use titanium-based alloys, aluminum-based alloys, or more generally, stainless steel alloys with a base of steel, nickel and/or chromium. 
       FIG. 4  shows an exploded view of several mechanical parts which, when assembled, form the probe  25 . 
     A body  47  forms the base  38  and the enclosures  41  and  42 . The channel  39  can be made directly in the body  47  by additive manufacturing. 
     The body  47  can remain open at its trailing edge, for example to arrange, in the body, the pneumatic channels making it possible to measure the total pressure. 
     Alternatively, these channels may also be made using the additive manufacturing method. 
     The trailing edge  31   a  of the mast  31  and the end  36  of the Pitot tube can be closed using a stopper  48  that can be made using any type of manufacturing method. 
     The shapes of the stopper  48  are simpler than those of the body  47 . It is for example possible to produce the stopper  48  by molding. 
     The additive manufacturing can of course also be used for the stopper  48 . 
     A support  49  can complete the probe  25 . 
     The support  49  can be used to support the pneumatic connector in a first part  49   a  as well as the evaporator  14  in a second part  49   b.    
     The support is assembled to the body  47  by the base  38 . 
       FIGS. 5 a  and 5 b    show another aerodynamic probe  60 . 
     More specifically, the probe  60  forms a piece of aeronautic equipment comprising a part  61  intended to be flush with the skin  27  of the aircraft. 
       FIG. 5 a    is a view in the plane of the skin  27  near the probe  60 . 
       FIG. 5 b    is a sectional view perpendicular to the plane of the skin  27 . 
     The part  61  is for example in the shape of a disc closing off an orifice  62  of the skin  27 . The orifice  62  is provided to receive the part  61  that is fastened by screwing on the skin  27 . 
     The probe  60  is for example a static pressure probe having one or more pressure taps  63  formed from channels emerging substantially perpendicular to the skin  27 . 
     The channel  39  circulates in the part  61 . The channel winds around pressure taps  63  in order to heat the part  61  and prevent the pressure taps from being closed off [by] ice. 
     In this embodiment, the channel  39  can also form a single loop or several parallel loops outside the evaporator  14 . 
     The probe  60  also comprises a part  65  inside the skin  27 . The inner part  65  makes it possible to receive a pressure sensor connected to the pressure taps in order 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 . 
     Like for the probe  25 , the part  61  is advantageously made by carrying out an additive manufacturing method. 
     The invention also relates to a data file stored on storage means and able to be loaded in the memory of a processing unit associated with an additive manufacturing machine able to manufacture an object by superimposing layers of material, which comprises three-dimensional depiction data of the piece of equipment as previously described, so as to allow, when it is loaded in the memory of, and processed by, said processing unit, the manufacture of said piece of equipment by said additive manufacturing machine. 
       FIG. 6  shows a probe  25  fastened to the skin  27  of an aircraft. 
     In this example embodiment, the evaporator  14  of the heating means is associated with functional means  80  of a second functional piece of equipment  81  of the aircraft forming a heat source for example giving off heat through loss of the heat during their operation. 
     Thus for example, the functional means of the second functional piece of equipment  81  of the aircraft can be formed by a part of or an electronic board thereof making it possible to contribute heat energy to the evaporator, as illustrated in this  FIG. 6 . 
     Of course, still other embodiments may be considered, and these means may be formed by an actuator of the second functional piece of equipment  81  of the aircraft or a part of or a piece of electrical equipment thereof. 
     Still other embodiments may be considered.