• 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 an artificial satellite comprising: - a support structure carrying equipment-carrying walls, - a launcher adapter integral with the support structure, - a first radiator (26), - at least a first transport equipment fluid heat (42) having at least one pipe (44) having a first heat exchange section (50) and a second heat exchange section (52), said second heat exchange section (52) being adapted to being in thermal contact with said first radiator (26), characterized in that said first heat exchange section (50) is in thermal contact with at least a portion of the starter adapter (16). The invention also relates to a method of filling a propellant gas tank of said artificial satellite.
    公开号:FR3031969A1
    申请号:FR1550613
    申请日:2015-01-27
    公开日:2016-07-29
    发明作者:Andrew Nicholas Walker
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    [0001] BACKGROUND OF THE INVENTION The invention relates to the technical field of artificial satellites, and more specifically to the misalignment of antenna reflectors and the storage of propellant gas. The known artificial satellites 1, diagrammatically represented in FIGS. 1, 2 and 4, generally consist of a support structure 14 carrying six planar carrier walls 6, 10 forming a box with cubic or parallelepipedal equipment. Antenna reflectors 18 are attached to the base of the equipment plane walls 6, 10 or directly on the support structure. An adapter-launcher 16 (commonly called by the acronym LVA of the English "Launch Vehicle Adapter") is fixed directly to the base of the support structure 14. This adapter-launcher is intended to be removably secured to an integral complementary adapter a satellite launcher. The launching adapter 16 is released from the complementary adapter of the launcher, when the satellite is dropped at the end of the propulsive phase of the launcher.
    [0002] The support structure 14 must at the same time be very light and resist launching by supporting several g of static acceleration. It is therefore generally made of carbon fiber. This material is very solid and has a coefficient of thermal expansion close to zero so that the support structure is deformed little. On the contrary, the adapter-launcher 16 is made of aluminum. This material is able to withstand a significant local load during launch. But, this material has a high coefficient of thermal expansion. However, in station, the adapter-launcher 16 periodically undergoes temperature variations of up to 60 ° C during its successive passages from the side of the sun 9 to the night side. As a result, the starter adapter 16 contracts at midnight in local solar time (see Figure 1) and expands at midday in local solar time (see Figure 2). Since the adapter-launcher 16 is fixed directly on the support structure 14, these expansions and contractions 3031969 2 cause deformations of the structure -support and the base of the planar equipment walls 6, 8. These deformations cause Periodic misalignment of the antenna reflectors 18. FIG. 3 shows the variation of the TA temperature of the starter adapter 16 over time as well as the misalignment of the antenna reflectors 18 resulting therefrom. To avoid these misalignments, it is possible to fix the antenna reflectors at a greater distance from the starter-adapter, for example, on a median part of the plane plane-equipment carriers. However, this positioning reduces the focal length of the antenna reflectors. This reduction can be constraining for their performance. To avoid these misalignments, some manufacturers have developed complex mechanisms for re-pointing antenna reflectors. These mechanisms include sensors capable of capturing the misalignments, a processing unit adapted to calculate the new orientation to be given to the antenna reflectors and actuators adapted to reorient the antenna reflectors. Nevertheless, these re-pointing mechanisms are not totally satisfactory because they have difficulties in capturing the misalignment due to the absence of a fixed reference system in the space. In addition, they are very expensive. The object of the present invention is to provide an artificial satellite whose antenna reflectors are not or little depointed and which do not have the disadvantages of existing mechanisms.
    [0003] Advantageously, the artificial satellite according to the invention does not reduce the focal length of the antenna reflectors. Advantageously, the artificial satellite according to the invention is less expensive than satellites equipped with a re-pointing mechanism and is more reliable over time.
    [0004] To this end, the subject of the invention is an artificial satellite having a longitudinal axis extending in the launching direction, said artificial satellite comprising: a support structure carrying equipment-carrying walls, an adapter an integral support member of the support structure, a first radiator, at least one first fluid heat transport equipment comprising at least a first heat exchange section and a second heat exchange section, said second heat exchange section; thermal exchange being in thermal contact with said first radiator, characterized in that said first heat exchange section is in thermal contact with at least a part of the adapter-launcher so as to allow heat exchange between the adapter-adapter launcher and the first radiator.
    [0005] According to particular embodiments, the artificial satellite comprises one or more of the following characteristics: it comprises a section integral with the starter adapter, said profile extending in a plane perpendicular to the longitudinal axis of the satellite and wherein said profile carries and is in direct thermal contact with at least a portion of the first heat exchange section. - An equipment wall, said anti-earth equipment wall is attached to the adapter-launcher via said profile. At least a portion of the first heat exchange section is in direct thermal contact with an equipment wall, said anti-earth equipment wall; said anti-earth equipment wall being attached to the starter adapter. - It comprises a section integral with the starter adapter, said profile extending in a plane perpendicular to the longitudinal axis of the satellite. - The first heat exchange section comprises several sections 30 fixed to each other so as to form a half-polygon or a polygon. The first heat exchange section has the shape of a semicircle or the shape of a circle. - The first heat transport equipment comprises an omega-shaped main pipe and at least one L-shaped auxiliary pipe; At least a portion of the main duct forming said first heat exchange section, at least a portion of the auxiliary duct forming said second heat exchange section. - It comprises a propellant gas tank and wherein the starter adapter is a cylindrical body; said propellant tank being arranged inside said cylindrical body; said tank being in thermal contact with the starter adapter. - The tank is mounted directly on the starter adapter. - It comprises at least a second fluid heat transport equipment; said second heat transport equipment including a first section and a second section; said second section being in thermal contact with said first radiator and said first section being in thermal contact with a set of electric accumulators. It comprises an adhesive strip made of a thermally conductive material; said adhesive strip being secured, on the one hand, to the first heat exchange section and, on the other hand, to at least a portion of said profile. - The first heat exchange section comprises at least two heat pipes arranged next to each other along a direction perpendicular to the longitudinal axis of the satellite. The first heat exchange section comprises at least two heat pipes arranged next to each other, along a direction parallel to the longitudinal axis of the satellite. The first heat transfer equipment comprises at least one heater in thermal contact with the second heat exchange section, and a control unit capable of controlling the ignition of said at least one heater.
    [0006] 3031969 5 Before launch, the satellite tank must be filled with propellant. When the propellant is a gas such as xenon, the filling operation of the tank causes an increase in the pressure of the gas in the tank. Increasing this pressure causes the temperature of the gas and the reservoir to increase. As a result, it is necessary to limit the filling rate. Thus, whenever a small amount of gas is introduced into the tank, it is necessary to wait for the tank to cool to an ambient temperature before a small amount of gas can be introduced again. This process must be repeated several times to completely fill the tank. The cooling time of the tank is important. The tank is usually installed inside the satellite and is thermally insulated from the outside. As a result, the cooling of the tank is extremely slow so that the process of filling the tank can take several days. To decrease the duration of this filling, it is possible to blow fresh air into the interior of the satellite to help cool the tank. However, the introduction of fresh air from a ventilation system can bring particles and contaminants into the satellite.
    [0007] The invention also relates to a method for filling a reservoir of the artificial satellite mentioned above. This method comprises the following steps: - cooling of the tank by the first radiator; and - filling the tank with propellant.
    [0008] Advantageously, according to the invention, there is a significant thermal coupling between the tank and the radiator. As a result, it is now possible to cool the tank by bringing cold air into the outside radiators. Thus no contaminant is introduced inside the satellite. The filling time of the tank can be significantly reduced. During this method of filling the tank, the heaters will of course be extinguished.
    [0009] The invention will be better understood on reading the description which follows, given solely by way of example and with reference to the figures in which: FIG. 1 is a schematic view of an artificial satellite according to FIG. state of the art when the starter adapter is positioned on the night side; - Figure 2 is a schematic view of an artificial satellite according to the state of the art when the starter adapter is positioned on the side of the sun; FIG. 3 is two curves representing, one the temperature of the adapter-launcher of a satellite of the state of the art as a function of time, the other, the misalignment of these antenna reflectors by function of time; FIG. 4 is an exploded perspective view of an artificial satellite according to a first embodiment of the invention; FIG. 5 is a sectional view from above of a lower part of the artificial satellite illustrated in FIG. 4; FIG. 6 is a section of the artificial satellite illustrated in FIG. 4, the sectional plane being illustrated in FIG. 5; FIG. 7 is a section similar to the view of FIG. 6 of an alternative embodiment of the artificial satellite illustrated in FIG. 4; FIG. 8 is a sectional view similar to FIG. 5 of a second embodiment of the artificial satellite according to the invention; Fig. 9 is a perspective view showing part of a third embodiment of the artificial satellite according to the invention; FIG. 10 is two curves representing, one the temperature of the launching adapter 16 of a satellite according to the invention as a function of time, the other the misalignment of the antenna reflectors as a function of time; FIG. 11 is an exploded perspective view of a fourth embodiment of the artificial satellite according to the invention in which a reservoir is mounted; and FIG. 12 is a diagram showing the steps of the method according to the invention. With reference to FIGS. 4 and 5, an artificial satellite 2 according to the first embodiment of the invention has a longitudinal axis A-A extending in the launching direction. It comprises equipment-carrying plane walls 4, 6, 8, 10, 12 suitable for carrying equipment, a support structure 14 carrying said equipment-carrying walls and an adapter-launcher 16 integral with the support structure 14.
    [0010] The equipment includes, in particular, antenna reflectors 18, electronic equipment (not shown), powered by solar panels 20 and at least one set of electric accumulators 24. The antenna reflectors 18 are attached to the base of the walls equipment carrier 6, 10 of the east and west faces. Solar panels 20 are mounted on the equipment-carrying walls of the North 4 and South 8 faces. The equipment wall of the north face 4 is provided with a first radiator 26 and a second radiator 28. The door wall -Equipment of the South face 8 is equipped with a third radiator 30.
    [0011] The support structure 14 and the launcher adapter 16 may be constituted in different ways. Only an exemplary constitution shown in Figure 4 will be described in detail. The example of the support structure shown in FIG. 4 comprises a central cylinder 32 and four fins 34 fixed, on the one hand, to the central cylinder 32 and, on the other hand, to an equipment support wall 4, 25 6, 8, 10. The exemplary starter adapter 16 shown in FIGS. 4 to 7 has the shape of a ring. It is constituted by a cylindrical body 35 having a circular base. The center of the circular base is disposed on the longitudinal axis A-A of the satellite. The cylindrical body 35 is secured to the lower portion of the central cylinder 32 of the support structure. The equipment carrier wall 12 which extends perpendicular to the longitudinal axis AA of the launcher side, said anti-earth equipment carrier wall 12, is fixed to the launcher adapter 16 and to the central cylinder 32, by the A profile 36 is shown diagrammatically in FIGS. 4, 7, 8 and 9. The profile 36 has a square section. This bracket has a first branch 38 attached to the adapter-launcher and the central cylinder and a second branch 40 attached to the wall-equipment anti-land 12, for example by bolts. The first leg 38 encloses the starter adapter 16. The second leg 40 forms a flange extending outwardly in a plane perpendicular to the longitudinal axis A-A of the satellite. The adapter-launcher is generally made of aluminum or an aluminum-based alloy. The support structure 14 is generally made of carbon fiber. With reference to FIG. 5, the artificial satellite 2 comprises a first fluid heat transport equipment 42 capable of cooling or heating the starter adapter 16 to attenuate the temperature variations to which it is subjected.
    [0012] The first fluid heat transport equipment 42 is a two-phase equipment such as a heat pipe. It forms a heat exchange loop. It comprises a pipe 44 containing a heat transfer fluid or several pipes 44, 45 containing heat transfer fluids, said pipes being in thermal contact in pairs. It also includes a heater 46 and a control unit 48 for controlling the ignition of the heater. The pipe 44 comprises a first heat exchange pipe section 50 in thermal contact with the starter adapter 16 and a second heat exchange pipe section 52.
    [0013] The first heat exchange section 50 is constituted by a plurality of rectilinear pipe sections 54 fixed to each other so as to form a polygon 56. This polygon 56 is disposed as close as possible to the launcher adapter 16. The second section heat exchanger 52 is in thermal contact 30 with the first radiator 26 and with the heater 46. The first radiator 26 constitutes a heat sink or cold source. The heater 46 is a hot source. The heater may for example be located on the inner face of the radiator, on the pipes, or directly 3031969 9 at the adapter. The first radiator-heater assembly is therefore a cold or hot thermal source. Advantageously, the second heat exchange section 52 is fixed on one face of the satellite having a different orientation of the orientation of the face which carries the starter adapter. As a result, the second heat exchange section 50 is generally not exposed at the same time, at the same time as the starter adapter. Figures 6 and 7 illustrate two possible modes of arrangement of the first heat exchange section 50.
    [0014] With reference to FIG. 6, the first heat exchange section 50 consists of two lines 44, 45 and, in particular, two heat pipes. The two heat pipes 44, 45 are stacked one above the other, along a direction parallel to the longitudinal axis A-A of the satellite. The lower heat pipe 44 is in direct physical contact with the profile 36.
    [0015] With reference to FIG. 7, the first heat exchange section 50 consists of two lines 44, 45 and, in particular, two heat pipes. The two heat pipes 44, 45 are arranged next to each other, along a direction perpendicular to the longitudinal axis A-A of the satellite. In this second mode of arrangement, the two heat pipes 44, 45 are in direct physical contact with the equipment-earth-bearing wall 12. Preferably, a thermal adhesive tape 58 (in English "thermal strap"), made of a material thermally conductive, such as aluminum or an aluminum alloy, is disposed on the two heat pipes 44, 45 and on the first branch 38 of the profile. It realizes a thermal bridge between them. It improves the heat conduction between the first heat exchange section 50 and the starter adapter 16. The first heat transport equipment 42 is used to cool or heat the starter adapter 16, either passively in using the "natural" orientation of the artificial satellite 2 with respect to the sun, and in particular the orientation of the starter adapter 16 and the first radiator 26 with respect to the sun, either actively, for example by using the heater 46.
    [0016] According to a second embodiment of the invention shown in FIG. 8, the artificial satellite 2 further comprises a second 70 and a third 71 heat transport equipment 70, each suitable for cooling a set of accumulators. 24, during their operation. The second 70 and the third 71 heat transport equipment each comprise a pipe having a first section 74 and a second section 76. The second section of the second heat transport equipment is in thermal contact with the first radiator 26. The The second section 76 of the third heat transport equipment is in thermal contact with the third radiator 30. The artificial satellite 2 illustrated in FIG. 8 comprises the same technical elements as the artificial satellite 2 shown in the other figures. These technical elements have been referenced by the same numerical references and have not described a second time. The heater 46 and the control unit 48 have not been shown in FIG. 8 for reasons of simplification. The embodiment of FIG. 8 makes it possible to use the first radiator 26 and the third radiator 30 to cool both the adapter-launcher 16 / reservoir 39 and two sets of electric accumulators 24. This configuration is advantageous. especially for some telecommunication satellites having an adapter-launcher 16 oriented anti-earth. Indeed, the adapter-launcher 16 of these satellites is exposed to maximum sunlight at midday in local solar time. And since the electric accumulator assemblies 24 dissipate heat only when they are used, ie when the satellite is eclipsed, at midnight in local solar time, the same radiator 26, 30 can dissipate the heat of the electric accumulator assemblies at midnight and the heat of the adapter-launcher 16 at midday. Thus, a single radiator 26, 30 can be used for two functions: the cooling of the electric accumulator assemblies and the cooling of the adapter-launcher. Reheating needs are also reduced. The heater may be common for control of the starter adapter and accumulators. This configuration is also advantageous for the artificial satellites 2 having an adapter-launcher 16 arranged facing east or west. Advantageously, this configuration makes it possible to increase the performance of the radiators by providing a North-South thermal link between the first radiator 26 and the third radiator 30. According to a third embodiment of the artificial satellite 2 according to the invention, represented on FIG. 9, the first heat transport equipment 42 comprises two omega-shaped main ducts 80, 81 and four L-shaped auxiliary ducts 82, 83, 84, 85. The omega form consists of one half circle connected to two linear branches. The semicircle 86 of a main line 80 is in thermal contact with a semi-cylindrical portion of the starter adapter 16. A first auxiliary line 82 is in thermal contact with a linear branch of the main line 80 and with the first radiator 26. A second auxiliary pipe 83 is in thermal contact with the other linear branch of the main pipe 80 and with the third radiator 30. The semicircle 87 of the other main pipe 81 is in thermal contact with a complementary semi-cylindrical portion of the starter adapter 16. A third auxiliary line 84 is in thermal contact with a linear leg of the main line 81 and the first radiator 26. A fourth auxiliary line 85 is in thermal contact with the other linear branch of the main pipe 81 and with the third radiator 30. The semicircles 86, 87 are arranged on the profile 36 or on the wall 25 equipment carrier 12. They are in thermal contact with the adapter-launcher 16. They constitute the first heat exchange section 50. A portion of each auxiliary pipe 82, 83, 84, 84 constitutes the second exchange section As a variant, the semicircles 86, 87 are replaced by half-polygons. Alternatively, the main pipe in the shape of omega (or half-polygon) is made in one piece, that is to say using a single conduit 3031969 12. This variant is advantageous for integration since it involves only 3 flat parts which facilitates their integration. In this embodiment, there is a north-south thermal link between the first radiator 26 and the third radiator 30, which is advantageous in terms of radiator performance. This embodiment can be easily combined with the previous embodiment: the first radiator 26 and the third radiator 30 may be common with the battery radiator. In operation, when the starter adapter 16 is oriented towards the sun at 12:00 in local solar time, the starter adapter 16 is a hot source for the first heat transport equipment 42. The heat recovered by the first section heat exchange 50 is evacuated by the first radiator 26 which is in thermal contact with the second heat exchange section 52. The first radiator 26 cools the adapter-launcher 16. On the contrary, when the adapter-launcher 16 is oriented on the shadow side, at 15 Oh00 in local solar time, the starter adapter 16 is heated by the heater 46 via the first heat transport equipment 42. Thus, the first heat transport equipment 42 can cool the adapter-launcher 16, when it has a high temperature, and warm it when it has a low temperature. Thus, as shown in Fig. 10, the temperature variation range of the starter adapter 16 is reduced so that the starter adapter does not deform or little. The misalignment has antenna reflectors 18 are thus reduced or even eliminated. An example of heat exchange value between the starter adapter and the first radiator can be estimated by the calculation described below.
    [0017] The objective is, for example, to limit the temperature variations of the adapter-launcher at 20 ° C +/- 10 ° C. Thus, the maximum temperature of the starter adapter shall not be greater than 30 ° C (303 K) when the starter adapter is in the sun, and shall not be less than 10 ° C (283 K) when the 'launcher adapter is in the shadows. The exposed portion of the cylindrical body 35 may be likened to a ring of 1.66 meters in outer diameter and 1.56 meters in inner diameter. The area exposed to the sun (Surf.) Is 0.254 m2. The surface coating is generally a hard coating, for example a hard anodizing coating such as alocrom brand, with a solar absorptivity (as) of about 0.5 and an infra-red emissivity (cIR) of about 0.1. When the starter adapter is exposed to the sun, it absorbs the heat Q ', Q, 01 = F, 01x as x Surf.
    [0018] C = 1.427 W / m 2 x 0.5 x 0.254 m 2 = 181 W. In which F 'is the mean solar flux. A part Qrad of this heat Q ', is directly radiated in the space Qrad.
    [0019] Qrad = EIR x Surf. x cy x T4 Qrad = = 12 W. In which 6 is the Stefan-Boltzmann constant And T is the temperature The amount of heat of heat to be exhausted by the heat pipe system is then Q ', o - Qrad = 169 W When the starter adapter is in the shade, the amount of heat from the sun and absorbed by the starter adapter is zero, the cylindrical body 35 radiates less energy to the space (because the temperature is lower ). In this case, Qrad = = EIR x Surf. x cy x T4 Qrad = 9VV. The first radiator 26 and heater assembly can be controlled at a constant temperature of typically 13 ° C. The first radiator is sized to reject the amount of heat when the starter adapter is fully exposed to the sun, the control unit and the heater can then be implemented to gradually provide heat as it goes. that the adapter-launcher goes into the shadows. In this case, the temperature difference between the starter adapter and the first radiator is 17 ° C when the starter adapter is totally exposed to solar radiation, and -3 ° C when the starter adapter is located in the shade.
    [0020] The minimum required thermal coupling is then determined by the amount of heat of heat to be discharged divided by the temperature gradient. When the starter adapter is exposed to direct sunlight, the thermal coupling is -169W / 17 ° C = 10 VV / K. When the starter adapter is in the shade, the thermal coupling is - 12W! 3 = 4 VV / K. Thus, a thermal coupling of about 10 W / K between the adapter-launcher and the first radiator 26 is necessary to reach a temperature of 20 ° C +/- 10 ° C on the cylindrical body 35 of the adapter-launcher . Such a value can easily be obtained by using heat pipes or other fluid heat transport systems. This calculation is indicative - the exact value will depend on the actual size of the cylindrical body of the adapter-launcher, and the thermo-optical coating installed on the surface of the cylindrical body of the adapter-launcher.
    [0021] An artificial satellite 2 according to a fourth embodiment of the invention is similar to the artificial satellite according to the first embodiment except that the cylindrical body 35 of the starter adapter comprises a succession of orifices 37 intended for fixing a reservoir 39 containing a propellant gas. The succession of orifices 37 extend along a circle contained in a plane perpendicular to the longitudinal axis A-A. The reservoir 39 is mounted inside the lower part of the housing created by the central cylinder 32. It is fixed to the starter adapter 16 via a fixing skirt 60 which encircles the tank. The lower part of the fixing skirt is provided with a serration 62 provided with orifices 64. The reservoir 39 is, for example, fixed to the launcher adapter 16 by nuts mounted in the orifices 64 of the skirt and fixing in the orifices 37 of the adapter-launcher 16. The reservoir 39 contains a gas or a propellant liquid or propellant, such as for example xenon or a gas containing xenon. From the mechanical point of view, it is advantageous to install the tank as low as possible in the satellite: this makes it possible to reduce the height of the center of gravity and thus makes it possible to reduce the forces at launch, which allows a lighter structure and less expensive. The ideal is to mount the tank on the starter adapter.
    [0022] Advantageously, the cooling of the adapter-launcher 16 makes it possible to mount the reservoir 39 therein. However, the amount of propellant that can be stored in the tank 39 is dependent on the temperature of that propellant. For example, if one considers the example of a reservoir having a fixed volume V ', and containing an ideal gas satisfying the equation of state pV = nRT, the quantity of gas n which can be stored is then equal at n = pV / RT. However, the pressure is limited by the volume of the tank at a maximum pressure p, above which there is a risk of explosion. Thus, the maximum amount of gas that can be stored is given by the following equation: ## EQU1 ## This amount is inversely proportional to the maximum temperature of the reservoir. Accordingly, if the tank is exposed to high temperatures, then the amount of gas stored in the tank must be reduced so as to remain below the maximum allowable pressure pmax. In this example, when the propellant is a perfect gas, a temperature reduction of 60 ° C (333K) at 20 ° C (293K) increases the amount of gas stored by 13%. However, propellants used in satellites are generally not ideal gases and the reduction in propellant temperature makes it possible to increase the amount of gas stored more substantially. For example, when xenon is used as propellant, the tank has a volume of 1 m 3 and the maximum pressure pm 'is 100 bar. At a temperature of 60 ° C, the maximum mass of xenon that can be injected into the reservoir would be 900 kg, while at a temperature of 20 ° C, the maximum mass of Xenon that can be injected is 1850 kg. . As a result, the increase in introduced xenon mass is greater than 100%. In addition, some liquid propellants are not compatible with high temperatures because of chemical degradation problems.
    [0023] Since the reservoir 39 is mounted within the starter adapter 16, there is significant thermal coupling therebetween so that the reservoir will tend to have the same temperature as the starter adapter 16. In this case, Indeed, even if an attempt is made to thermally isolate the tank from the starter adapter, the support structure of the tank 64 must be rigidly connected to the starter adapter and the tank, thereby creating a thermal connection. The mechanical requirements in practice make it difficult to have a thermal coupling of less than 0.5 W / K between the tank and the adapter-launcher.
    [0024] In any case, the effect of such isolation is only to slow down the transient thermal response of the reservoir. If the starter adapter 16 is hot for several days, this isolation has a minimal effect on the reservoir in steady state at a steady temperature. During certain phases of the mission, the starter adapter 16 is constantly illuminated by the sun for several days, and becomes very hot - generally 70 ° C. In these cases, the temperature of the tank tends to be identical to the temperature. of the adapter-launcher. Such a temperature is too high for the storage of propellant. It is then not possible to mount a tank directly on the adapter-launcher 16.
    [0025] The invention makes it possible to install a reservoir directly on the central cylinder 32 of the launching adapter 16. The maximum temperature of the central cylinder 32 of the adapter-launcher is brought to a lower level, and consequently the stabilized temperature The reservoir is reduced to within acceptable limits for the storage of the propellant. The temperature of the tank remains acceptable even if the thermal coupling between the tank and the starter adapter 16 is increased. Thus, the first heat transport equipment 42 is also used to cool and control the temperature of the reservoir 39. Finally, the first heat transport equipment 42 is advantageously used to cool the reservoir 39 as it is being filled. For this purpose, during a step 66 visible in FIG. 10, the first radiator 26 cools the reservoir 39 via the launching adapter 16. Then, during a step 68, the reservoir 39 is filled with gas. During this filling step 68, the first radiator 26 continues to cool the reservoir 39 to increase the filling capacity of the reservoir 39 or to reduce the filling time thereof. Steps 66 and 68 may be repeated several times.
    [0026] Advantageously, the cooling of the second heat exchange section 52 makes it possible to cool the starter adapter 16 and, by contact, the reservoir 39 fixed thereto. The cooling of the reservoir 39 makes it possible to cool the gas injected inside the reservoir. This cooling is opposed to the natural warming of the gas due to its compression during the injection. As indicated above, the temperature of the propellant in flight does not depend much on the thermal coupling between the tank and the starter adapter. The mechanical installation of the tank can then be designed with a strong thermal coupling so as to maximize the cooling effect during filling. As indicated above, this thermal coupling is favored if the tank is mounted directly on the launcher interface; that is to say without intermediate piece. Thus, the filling of the reservoir 39 is more secure and can be achieved more quickly. In addition, the cooled reservoir 39 can store a larger amount of gas for the reasons mentioned above. Alternatively, the profile is formed of a block with the adapter-launcher. In a variant also, the profile extends towards the inside of the cylinder of the adapter-launcher. In this case, the tank is not mounted in the starter adapter. Alternatively, the radiator in contact with the heat transport equipment or 42, 70, 71 is disposed on the south face, the east face, the west face, the earth face or the anti-earth face of the satellite. Alternatively, the north equipment wall has a single radiator 25 which can be used for both the first 70 and the second 71 heat transport equipment. Alternatively, the first heat exchange section is circular. Alternatively, it may be envisaged to combine the different possibilities of the two modes of arrangement illustrated in FIGS. 6 and 7. For example, as an alternative to the first arrangement, it is possible to arrange the two heat pipes 44, 45, one next to the other, on and in direct contact with the profile 36. Similarly, it is possible to arrange an adhesive tape 58 on the two heat pipes 44, 45 stacked and on the first branch 38 of the profile. It can also be envisaged to arrange the two heat pipes 44, 45 stacked on the earth-equipment carrier wall 12. In a variant, the second heat exchange section 52 of the pipe 44 is in thermal contact with the first heat sink 26. and the third radiator 30. Alternatively, the first fluid heat transport equipment 42 is a monophasic or triphasic equipment. Alternatively, the starter adapter may have a cylindrical shape with a square base or a rectangular base or polygonal base whose center is disposed on the longitudinal axis of the satellite. Alternatively, the support structure may have different shapes and in particular a frustoconical shape or a faceted bowl shape with a central chimney or consist of panels assembled in the form of a cross or even consist only of a wall plane door equipment extending perpendicularly to the longitudinal axis AA of the satellite.
    权利要求:
    Claims (16)
    [0001]
    1. Artificial satellite (2) having a longitudinal axis (AA) extending in the launching direction, said artificial satellite (2) comprising: a support structure (14) carrying equipment-carrying walls (4, 8, 6, 10, 12), - an adapter-launcher (16) integral with the support structure (14), - a first radiator (26), - at least one first fluid transport equipment (42) per fluid having at least a first heat exchange section (50) and a second heat exchange section (52), said second heat exchange section (52) being in thermal contact with said first heat sink (26), characterized in that said first heat exchange section (50) is in thermal contact with at least a portion of the starter adapter (16) so as to allow heat exchanges between the starter adapter (16) and the first radiator ( 26).
    [0002]
    2. artificial satellite (2) according to claim 1, which comprises a section (36) integral with the starter adapter (16), said section (36) extending in a plane perpendicular to the longitudinal axis (AA). ) of the satellite, and wherein said profile (36) carries and is in direct thermal contact with at least a portion of the first heat exchange section (50).
    [0003]
    3. Artificial satellite (2) according to claim 2, wherein a device-carrying wall, said anti-earth equipment wall (12), is fixed to the adapter-launcher (16) via said profile (36).
    [0004]
    4. Artificial satellite (2) according to claim 1, wherein at least a portion of the first heat exchange section (50) is in direct thermal contact with a wall equipment, said wall-equipment anti-land (12); said anti-earth equipment wall (12) being attached to the starter adapter (16). 303 196 9 20
    [0005]
    5. Artificial satellite (2) according to claim 4, which comprises a section (36) integral with the starter adapter (16), said section (36) extending in a plane perpendicular to the longitudinal axis (AA). ) of the satellite (2).
    [0006]
    An artificial satellite (2) according to any one of claims 1 to 5, wherein said first heat exchange section (50) comprises a plurality of sections (54) attached to each other so as to form a half-polygon or a polygon (56).
    [0007]
    7. An artificial satellite (2) according to any one of claims 1 to 5, wherein said first heat exchange section (50) has the shape of a semicircle (86, 87) or the form of a circle.
    [0008]
    An artificial satellite (2) according to any of claims 1 to 7, wherein the first heat transport equipment (42) comprises an omega-shaped main conduit (80, 81) and at least one auxiliary conduit (82, 83, 84, 85) L-shaped; at least a portion of the main pipe (80, 81) forming said first heat exchange section (50), at least a portion of the auxiliary pipe (82, 83, 84, 85) forming said second heat exchange section (52).
    [0009]
    9. artificial satellite (2) according to any one of claims 1 to 8, which comprises a reservoir (39) of propellant and wherein the starter adapter (16) is a cylindrical body (35); said tank (39) of propellant gas being arranged inside said cylindrical body (35); said reservoir (39) being in thermal contact with the starter adapter (16).
    [0010]
    10. An artificial satellite (2) according to claim 9, wherein the reservoir (39) is mounted directly on the starter adapter (16).
    [0011]
    11. An artificial satellite (2) according to any one of claims 1 to 10, which comprises at least one second heat transport equipment (70, 71) per fluid; said second heat transport equipment (70, 71) comprising a first section (74) and a second section (76); said second section (76) being in thermal contact with said first radiator (26) and said first section (74) being in thermal contact with a set of electric accumulators (24).
    [0012]
    12. Artificial satellite (2) according to any one of claims 2, 3, 5 to 11, which comprises an adhesive strip (58) made of a thermally conductive material; said adhesive strip (58) being secured, on the one hand, to the first heat exchange section (50) and, on the other hand, to at least a portion of said profile (36).
    [0013]
    13. Artificial satellite (2) according to any one of claims 1 to 12, wherein the first heat exchange section (50) comprises at least two heat pipes (44, 45) arranged, one next to the another, along a direction perpendicular to the longitudinal axis (AA) of the satellite.
    [0014]
    14. Artificial satellite (2) according to any one of claims 1 to 12, wherein the first heat exchange section (50) comprises at least two heat pipes (44, 45) arranged next to the other, along a direction parallel to the longitudinal axis (AA) of the satellite.
    [0015]
    15. Artificial satellite (2) according to any one of claims 1 to 14, wherein the first heat transfer equipment (42) comprises at least one heater (46) in thermal contact with the second heat exchange section (52), and a control unit (48) adapted to control the ignition of said at least one heater (46).
    [0016]
    16. A method of filling a reservoir (39) of an artificial satellite (2) according to claim 1; characterized in that it comprises the following steps: - cooling (66) of the reservoir (39) by the first radiator (26); and filling (68) the tank (39) with propellant.
    类似技术:
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    同族专利:
    公开号 | 公开日
    EP3212503A1|2017-09-06|
    WO2016120547A1|2016-08-04|
    CN107548370A|2018-01-05|
    EP3212503B1|2017-11-22|
    JP6334827B2|2018-05-30|
    JP2018502773A|2018-02-01|
    CN107548370B|2019-05-07|
    US9902507B2|2018-02-27|
    FR3031969B1|2017-01-27|
    US20170361951A1|2017-12-21|
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    法律状态:
    2016-01-28| PLFP| Fee payment|Year of fee payment: 2 |
    2016-07-29| PLSC| Search report ready|Effective date: 20160729 |
    2017-01-26| PLFP| Fee payment|Year of fee payment: 3 |
    2018-01-29| PLFP| Fee payment|Year of fee payment: 4 |
    2019-09-27| ST| Notification of lapse|Effective date: 20190906 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1550613A|FR3031969B1|2015-01-27|2015-01-27|ARTIFICIAL SATELLITE AND METHOD FOR FILLING A PROPULSIVE GAS TANK OF SAID ARTIFICIAL SATELLITE|FR1550613A| FR3031969B1|2015-01-27|2015-01-27|ARTIFICIAL SATELLITE AND METHOD FOR FILLING A PROPULSIVE GAS TANK OF SAID ARTIFICIAL SATELLITE|
    PCT/FR2016/050136| WO2016120547A1|2015-01-27|2016-01-22|Artificial satellite and method for filling a tank of propellent gas of said artificial satellite|
    EP16705226.5A| EP3212503B1|2015-01-27|2016-01-22|Artificial satellite and method for filling a tank of propellent gas of said artificial satellite|
    US15/546,118| US9902507B2|2015-01-27|2016-01-22|Artificial satellite and method for filling a tank of propellent gas of said artificial satellite|
    CN201680009300.3A| CN107548370B|2015-01-27|2016-01-22|Artificial satellite|
    JP2017539323A| JP6334827B2|2015-01-27|2016-01-22|Artificial satellite and method for filling a propellant gas tank of the artificial satellite|
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