JP2004196050A - Artificial satellite - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To provide a high-speed flying object such as a small rocket for launch of a pico type or micro type artificial satellite, a fairing included in the head of the high-speed flying object, and a manufacturing method for the fairing. <P>SOLUTION: The artificial satellite 2 is formed integrally with the fairing 6 used as a shell, and the artificial satellite 2 is mounted to the upper part of a first stage solid motor 3 via a second stage solid motor 4 and an inter-stage section 5, thereby constituting a two-stage type rocket 1. The fairing 6 is provided with an ablator layer as a thermal protection system, and the surface of the ablator layer is coated with heat resistant paint having low sunlight absorptance to form a low sunlight absorptance layer. A plurality of ventilation holes are formed in the low sunlight absorptance layer, and pyrolysis gas generated from the ablator layer by aerodynamic heating provided at the flying time of the two-stage type rocket 1 is discharged out of the the fairing 6 through the ventilation holes. <P>COPYRIGHT: (C)2004,JPO&NCIPI

Description

[0001]
TECHNICAL FIELD OF THE INVENTION
The present invention relates to an artificial satellite that travels in orbit where a rare atmosphere exists.
[0002]
[Prior art]
The main operational orbit of a satellite is divided into a low orbit at an altitude of several hundred km and a geosynchronous orbit at an altitude of about 36000 km. Of these, the atmosphere is present in the low orbit, albeit sparse, and the artificial satellite traveling in the low orbit receives air resistance during its travel, and its traveling speed decreases. A satellite operated in such a low orbit must travel at a speed corresponding to its altitude to obtain centrifugal force that balances the gravity from the earth, and if it is decelerated by air resistance, The altitude of the satellite will decrease. Then, it is necessary to increase the speed of the artificial satellite by such a deceleration amount and to raise the artificial satellite by an amount corresponding to a decrease in altitude. . Therefore, if the satellite is susceptible to large air resistance, the deceleration and altitude decrease as described above will increase, and a large amount of fuel will be required to recover the deceleration and altitude decrease. This leads to weight and weight.
[0003]
Therefore, Japanese Patent Application Laid-Open No. 2000-21596 discloses an artificial satellite in which the cross-sectional area of the satellite main body in the traveling direction is made small and elongated so as to be able to accommodate on-board equipment (Patent Document 1). Such an artificial satellite has a hexahedral shape that is elongated in the traveling direction, and is configured to have the same cross-sectional shape over the entire length in the longitudinal direction in order to secure a space for accommodating on-board equipment. This makes it possible to relatively reduce the air resistance that the artificial satellite receives when traveling in orbit.
[0004]
[Patent Document 1]
JP-A-2000-21596
[0005]
[Problems to be solved by the invention]
However, in the artificial satellite disclosed in Japanese Patent Application Laid-Open No. 2000-21596, since the cross-sectional shape and the cross-sectional area of the tip portion are the same as the cross-sectional shape and the cross-sectional area of the other portions, a large amount of air is generated at the tip surface. This will cause resistance, which will cause deceleration.
[0006]
The satellite disclosed in the above publication is housed in a fairing provided on the head of a rocket in order to protect the satellite from air resistance received during the launch at the time of the launch. Because the shape of the disclosed satellite is limited, the fairing designed according to this shape becomes large, and in some cases, the rocket itself must use an unnecessarily large one. there were. Furthermore, if the satellite has an extremely thin shape, there may be a lot of extra space when the satellite is stored in the fairing, and it must be launched using an unnecessarily large rocket. Sometimes.
[0007]
Satellites whose mass is 1000 kg or more are classified as large, 500 kg to 1000 kg are small, 100 kg to 500 kg are mini, 10 kg to 100 kg are micro, and 10 kg or less are pico. Sometimes. Among such satellites, large, small, and mini-type satellites are stored in a fairing provided on the head of a rocket dedicated to launching the satellite, and the rocket uses the rocket at a desired time and at a desired altitude and altitude. It is transported to orbit, the fairing is discarded, and put into this orbit.
[0008]
Once satellites have begun operation in orbit, they should be easily repaired and refurbished like ground equipment, even if a problem occurs that makes it impossible to carry out the scheduled mission. Can not. For this reason, the equipment and components for artificial satellites are subjected to a thorough test and inspection before launch so as to prevent inconvenience as much as possible in orbit, and their reliability is ensured.
[0009]
In particular, large and small satellites have complicated systems, a wide range of missions, and a large number of equipment and components, so many tests and inspections to ensure reliability are conducted over a long period of time. Will be done. In addition, many tests and inspections are required for devices used for assembling such artificial satellites.
[0010]
On the other hand, pico-type and micro-type satellites have relatively simple systems, many of which are dedicated to a single mission, and have a relatively small number of equipment and parts, which makes them large and small. There is an advantage that the number and the period of tests and inspections required are smaller than those of the artificial satellite, and the development period and the development cost are reduced.
[0011]
In addition, earth observation satellites and communication satellites are required to reduce the development period as much as possible in order to respond to sudden changes in the global environment and obsolescence of satellite-borne equipment. There is an increasing demand for pico-type and micro-type satellites for use in applications.
[0012]
Such pico-type and micro-type artificial satellites are mounted on a launch vehicle for launching another large or small artificial satellite by a so-called piggyback method and launched. This piggyback launch method involves mounting a pico-type or micro-type satellite in the extra space inside the fairing where other large or small satellites are stored, and attaching the large or small satellite to the rocket body. When the satellite is departed from the satellite, a pico-type or micro-type artificial satellite is also departed.
[0013]
However, when the artificial satellite disclosed in the above publication is configured as a pico-type or micro-type artificial satellite, the extra space inside the fairing has many restrictions on the space shape, and is disclosed in the above publication. Due to the shape limitations of the satellites used, there are cases where the satellite cannot be mounted on a desired rocket.
[0014]
In the case of a piggyback type rocket, the launch of a pico-type or micro-type satellite must be synchronized with the launch of a large or small satellite, which is the main purpose of launching the rocket. Even if a satellite can be developed in a short period of time, it is often not possible to launch it early, and the benefits of shortening the development period cannot be fully enjoyed.
[0015]
In the case of a piggyback type rocket, when a large or small satellite is separated from the rocket body, the pico-type or micro-type satellite is also separated. In some cases, satellites could not be put into optimal altitude and orbit.
[0016]
The present invention has been made in view of such circumstances, and reduces the air resistance received by an artificial satellite during orbital operation as much as possible, and at the same time reduces the spatial efficiency when mounted on a rocket used for launch. An object of the present invention is to provide an artificial satellite which is significantly improved as compared with the above.
[0017]
[Means for Solving the Problems]
In order to solve the above problems, the artificial satellite according to the present invention is an artificial satellite that travels in orbit in which a rare atmosphere exists, in which the outer shell has a streamline shape tapered in the traveling direction. It is characterized by functioning as a fairing at launch.
[0018]
By adopting such a configuration, the cross-sectional area of the leading end of the satellite in the traveling direction is significantly reduced as compared with the conventional one, and the air resistance received by the satellite in orbit is reduced as much as possible. it can.
[0019]
Further, since the outer shell of the artificial satellite is used as a fairing, it is not necessary to provide a fairing separately from the artificial satellite, and it is possible to launch using a rocket having a size optimal for the artificial satellite. Therefore, in the artificial satellite according to the present invention, the spatial efficiency when the artificial satellite is mounted on the rocket used for launch can be greatly improved as compared with the conventional one, and as a result, the rocket for launching the artificial satellite can be improved. A drastic reduction in size and cost can be realized as a whole.
[0020]
Further, in the artificial satellite according to the present invention, since the outer shell structure of the artificial satellite and the strength member of the outer shell also serve as the fairing, compared with the total mass of the conventional artificial satellite and the fairing. Therefore, the mass of the satellite can be significantly reduced. This makes it possible to significantly reduce the size and weight of the launch rocket as compared to conventional satellites.
[0021]
In addition, by configuring such a small and lightweight launch vehicle, the amount of launch fuel of the launch vehicle can be significantly reduced.
[0022]
In addition, since the launch vehicle can be configured to be small and lightweight, the time, labor, cost, and the like involved in the development, manufacture, and launch of the rocket can be significantly reduced as compared with the conventional technology, and therefore, It can be used exclusively for launching pico-type or micro-type satellites constructed by applying the invention.
[0023]
As described above, when the artificial satellite according to the present invention is configured as a pico-type or micro-type artificial satellite, a dedicated rocket for launching the artificial satellite can be configured, so that a pico-type or micro-type artificial A satellite can be launched, and a pico-type or micro-type satellite can be put into an optimal altitude and orbit.
[0024]
Further, in the above invention, it is preferable that the outer shell has a heat-resistant structure against aerodynamic heating.
[0025]
As described above, the rocket for launching a satellite according to the present invention is significantly smaller and lighter than the conventional rocket, so that the flight speed is higher than that of a conventional multi-stage rocket having three or more stages for launching a satellite. Will be significantly faster and the fairing will be exposed to aerodynamic heating. Therefore, by using a structure in which the outer shell is provided with a heat-resistant structure as described above, even in a severe flying environment caused by high-speed flight, it is possible to ensure that the equipment and parts inside the outer shell are not damaged by aerodynamic heating without being damaged. Can be transported to altitudes.
[0026]
Further, in the above invention, the outer shell includes an ablator layer and a low solar absorptivity layer made of a heat-resistant film having a low solar absorptivity and covering the surface of the ablator layer. It is preferable that the solar absorptivity layer is provided with a vent through which the pyrolysis gas generated from the ablator layer passes to be released to the outside.
[0027]
Many rockets flying at high speed have an ablator as a thermal protection system. This ablator is configured as the outer wall of a rocket, etc. When the rocket flies, its outer surface is exposed to high temperatures, so that the synthetic resin that is the constituent material evaporates and the endothermic reaction at this time causes the inside of the ablator to ablate. The temperature of the rocket is prevented from rising, and the pyrolysis gas covers the ablator surface, thereby preventing the rocket from being further heated. When the fairing is provided with such an ablator, the surface of the ablator is carbonized by being exposed to a high temperature due to aerodynamic heating, and becomes black. There is no problem since the fairing mounted on a normal rocket is discarded in the air or in outer space, but in the satellite according to the present invention, the outer surface functions as a fairing, so that the surface of the ablator layer is In the configuration exposed to the outside of the outer shell, the solar heat absorption rate is high, causing an increase in the temperature inside the outer shell, which may cause failure of devices and components inside the outer shell.
[0028]
Therefore, by adopting the above configuration, the ablator layer not only protects the equipment and components inside the outer shell from aerodynamic heating at the time of flight, but also covers the entire head with the low solar absorptivity layer. Therefore, absorption of solar heat can be suppressed as much as possible, and as a result, it is possible to prevent equipment and components inside the outer shell from being damaged by heat as much as possible.
[0029]
In addition, since the vent hole is provided, the pyrolysis gas generated from the ablator layer passes through the vent hole and is released to the outside of the artificial satellite. It is possible to prevent the low solar absorptivity layer from peeling off due to stagnation inside.
[0030]
BEST MODE FOR CARRYING OUT THE INVENTION
Hereinafter, a two-stage rocket equipped with an artificial satellite according to an embodiment of the present invention will be specifically described with reference to the drawings. FIG. 1 is a side view showing an overall configuration of a two-stage rocket according to an embodiment of the present invention. As shown in FIG. 1, a two-stage rocket 1 according to an embodiment of the present invention includes an artificial satellite 2 according to the present invention, a first-stage solid motor 3, a second-stage solid motor 4, and an inter-stage portion 5. It is mainly composed of
[0031]
The artificial satellite 2 constitutes the head of the two-stage rocket 1 and has a generally cylindrical shroud portion 2a (see FIG. 2) and a rounded conical nose connected to the upper end of the shroud portion 2a. 2b (see FIG. 2). The mass of the artificial satellite 2 is, for example, 45 kg, and is classified as a micro-type artificial satellite. Below the artificial satellite 2, a first-stage solid motor 3 is connected via an interstage 5.
[0032]
The interstage 5 has a cylindrical shape having substantially the same diameter as the cylinder of the artificial satellite 2, and a second-stage solid motor 4 is supported inside.
[0033]
The first-stage solid motor 3 has a movable nozzle, and performs flight control by changing the injection direction of the nozzle. The second-stage solid motor 4 does not have a structure for performing special flight control.
[0034]
The launching process of the artificial satellite 2 is as follows. First, by starting combustion of the first-stage solid motor 3, the two-stage rocket 1 starts to rise from a launch field near the pole. At this time, the angle between the axis of the two-stage rocket 1 and the horizon when the two-stage rocket 1 is launched is 90 °. That is, the two-stage rocket 1 is launched vertically upward. The first-stage solid motor 3 keeps burning for 40 seconds, and at the end of combustion, its flying altitude reaches about 45 km. Next, at the end of combustion of the first-stage solid motor 3, that is, 40 seconds after the start of launch, the first-stage solid motor 3 is disconnected as soon as possible, and combustion of the second-stage solid motor 4 is started. . The second-stage solid motor 4 continues to burn for 39 seconds, and the flight speed at the end of combustion reaches the maximum flight speed of about 7.7 km / s, and the flight altitude reaches an altitude of about 240 km. After the completion of the combustion of the second-stage solid motor 4, the second-stage solid motor 4 and the interstage 5 are separated from the artificial satellite 2, and the artificial satellite 2 is put into a ballistic orbit.
[0035]
Further, another example of the launching process of the artificial satellite 2 is as follows. First, by starting combustion of the first-stage solid motor 3, the two-stage rocket 1 starts to rise from a launch site near the equator. At this time, the angle between the axis of the two-stage rocket 1 and the horizon when the two-stage rocket 1 is launched is 90 °. That is, the two-stage rocket 1 is launched vertically upward.
[0036]
About 39 seconds after the start of combustion of the first-stage solid motor 3, the two-stage rocket 1 starts a gravity turn to the east at a pitch of about 2 ° / sec. The first-stage solid motor 3 finishes burning 40 seconds after the start of combustion, and is separated from the two-stage rocket 1 as soon as possible.
[0037]
About two seconds after the completion of the combustion of the first-stage solid motor 3, the second-stage solid motor 4 starts burning. The second-stage solid motor 4 finishes combustion approximately 39 seconds after the start of combustion, at which time the angle between the axis of the two-stage rocket 1 and the horizon is 0 °, and the speed reaches 8.1 km / s. . The altitude reached at this time is about 185 km, and the artificial satellite 2 is put into orbit here. If the weight of the artificial satellite 2 is further reduced, the reaching altitude can be further increased.
[0038]
Such a two-stage rocket 1 flies at a significantly higher speed than the flight speed of a conventional multistage rocket having three or more stages. Therefore, the flying environment such as the heating rate, dynamic pressure, and enthalpy that the two-stage rocket 1 receives during flight is much harsher than that of a multi-stage rocket having three or more stages during flight. According to calculations by the inventors of the present application, the maximum aerodynamic heating rate of the two-stage rocket 1 launched from a launch field near the pole and flying in ballistic flight is about 2.6 MW / m. ^Two Where the maximum dynamic pressure is about 13 atmospheres and the maximum acceleration is about 31G. These are each ten times or more the size of a large rocket such as the H2A rocket developed by the Space Development Agency of Japan, and how the flight environment of the two-stage rocket 1 according to the present embodiment is. You can see if it is harsh.
[0039]
In order to withstand such a severe flying environment, the artificial satellite 2 according to the present embodiment has a configuration as described below. FIG. 2 is a schematic cross-sectional view schematically illustrating a configuration of the artificial satellite 2 according to the embodiment of the present invention. As shown in FIG. 2, the artificial satellite 2 has a fairing (nose fairing) 6 as its outer shell structure. That is, the artificial satellite 2 according to the present embodiment is formed integrally with the fairing 6. The fairing 6 is not discarded before the artificial satellite 2 is put into orbit, but is used as an outer shell of the artificial satellite 2.
[0040]
The fairing 6 is composed of a plurality of layers, and has a structure in which a low solar absorptivity layer 7, an ablator layer 8, a heat insulating layer 9, and a metal sheet layer 10 are sequentially stacked from the outside. The low solar absorptivity layer 7 uses a white paint having sufficient resistance to aerodynamic heating during flight of the two-stage rocket 1. In this way, by covering almost the entire fairing 6 with the low solar absorptivity layer 7 having a low solar absorptivity, when the fairing 6 is exposed to sunlight, the absorption of solar heat is reduced. The temperature rise inside the fairing 6 can be suppressed as much as possible.
[0041]
Here, the white color means a material having relatively high brightness regardless of achromatic color or chromatic color. That is, the white system includes, for example, a slightly bluish white or the like. Further, in the present embodiment, the low solar absorptivity layer 7 is configured to be a white-based layer. However, the present invention is not limited to this, and when the fairing 6 is exposed to sunlight, The color of the low solar absorptivity layer 7 does not matter as long as the layer has a low solar absorptivity such that the internal space of the fairing 6 can be maintained at a temperature allowed by the equipment provided therein. Needless to say.
[0042]
The ablator layer 8 is made of a material such as CFRP used as a rocket ablator. In the present embodiment, it is assumed that the ablator layer 8 is made of CFRP. FIG. 3 is a cross-sectional view illustrating a configuration of the ablator layer 8 according to the embodiment of the present invention. As shown in FIG. 3, the ablator layer 8 is further divided into a heat protection layer 8a and a structural layer 8b. The heat protection layer 8a, which is the outer layer, is made of CFRP in which a large number of strip-shaped cloth bodies obtained by cutting carbon fiber woven into a cloth shape into strip shapes are embedded in a synthetic resin. The structure is such that thermal protection performance can be achieved and thermal stress can be kept low.
[0043]
On the other hand, the structural layer 8b, which is the inner layer of the ablator layer 8, is made of CFRP in which a plurality of carbon fibers are woven into a continuous cloth, and has a high specific strength even at high temperatures even if carbonized. And a structure capable of maintaining high specific rigidity. With such a structure, it is not necessary to provide a metal structural layer on the fairing 6, and it is possible to realize strength and rigidity comparable to a metal fairing while being lightweight.
[0044]
The heat insulating layer 9 is made of a heat insulating material made of an inorganic material. As shown in FIG. 2, the ablator layer 8 and the heat insulating layer 9 are configured such that they are thickest at the stagnation point 6 a, which is the tip of the fairing 6, and become thinner from here toward the rear. This is because, at the stagnation point portion 6a, the aerodynamic heating during the flight becomes the maximum and becomes lower toward the rear, so that the portion closer to the front end portion of the fairing 6 is required to have higher resistance to the aerodynamic heating. Further, by reducing the thicknesses of the ablator layer 8 and the heat insulating layer 9 toward the rear of the fairing 6, there is also an effect that the weight of the fairing 6 can be further reduced.
[0045]
An aluminum sheet is used for the metal sheet layer 10.
[0046]
Inside the fairing 6, equipment-mounted panels 11, 12, and 13 are provided. Each of these equipment-mounted panels 11 to 13 includes various sensors, an on-board computer, a storage device, and an artificial device such as an antenna. Satellite internal equipment 14 is mounted. These internal devices 14 perform radiant heat exchange with the metal sheet layer 10, whereby the temperature of the internal devices 14 is, for example, −15 ° C., which is the allowable temperature range of general electronic devices. Appropriately maintained in the range of 45 ° C.
[0047]
FIG. 4 is a graph showing changes in the surface temperature of the fairing 6, the temperature of the metal sheet layer 10, and the temperature of the internal device 14 during the flight of the two-stage rocket 1 according to the embodiment of the present invention. As shown in FIG. 4, the surface temperature of the fairing 6 reaches a maximum of about 1690 ° C. When the two-stage rocket 1 flies, the fairing 6 is exposed to such a high temperature, so that a pyrolysis gas is generated from the thermal protection layer 8a, and the temperature rise inside the fairing 6 due to the endothermic reaction at this time. This prevents the pyrolysis gas from covering the surface of the fairing 6, thereby preventing further heating of the satellite 2.
[0048]
Further, when the fairing 6 receives such aerodynamic heating, the thermal protection layer 8a is carbonized and changes color to black. Here, when the low solar absorptivity layer 7 is not provided, the surface of the fairing 6 changes color to black over the whole, and absorbs sunlight during the orbital operation of the artificial satellite 2, It is conceivable that the internal temperature of the artificial satellite 2 abnormally increases. In the fairing 6 according to the present embodiment, the low solar absorptivity layer 7 is laminated on the surface of the ablator layer 8, and as described above, the low solar absorptivity layer 7 has sufficient resistance to aerodynamic heating. Since the low-light-absorbing layer 7 does not disappear due to aerodynamic heating when the two-stage rocket 1 flies, the surface of the fairing 6 is kept white. For this reason, the solar absorptivity of the surface of the artificial satellite 2 can be reduced, and the rise in the internal temperature of the artificial satellite 2 can be suppressed. The heat-resistant paint does not have to be applied to some portions where the aerodynamic heating is expected to be higher than the upper limit temperature of the heat-resistant paint.
[0049]
FIG. 5 is a partial perspective view of the fairing 6 showing the configuration of the low solar absorptivity layer 7 according to the embodiment of the present invention, and FIG. 6 is a schematic partial cross-sectional view of the fairing 6. As shown in FIGS. 3, 5 and 6, the low solar absorptivity layer 7 is provided with a large number of air holes 15. That is, the heat protection layer 8a is exposed only in the portion of the ventilation hole 15 in the surface of the fairing 6. As described above, while the fairing 6 is under aerodynamic heating, the thermal protection layer 8a generates pyrolysis gas. As shown in FIG. 6, this pyrolysis gas passes through the vent hole 15 and is discharged to the outside of the artificial satellite 2. On the other hand, when the ventilation hole 15 is not provided in the low solar absorptivity layer 7, the escape of the pyrolysis gas generated from the thermal protection layer 8a by aerodynamic heating is eliminated, and the low solar absorptivity is low. Cracking or peeling will occur in layer 7. Therefore, by adopting a configuration in which the ventilation holes 15 are formed in the low solar absorptivity layer 7, the occurrence of cracks or peeling of the low solar absorptivity layer 7 can be suppressed.
[0050]
In other words, by configuring the fairing 6 as described above, it is possible to prevent heat generated by aerodynamic heating during the flight of the two-stage rocket 1 from invading the satellite 2 and to operate the satellite 2 in orbit. It is possible to release the heat generated from the internal device 14 to the outside of the aircraft during the flight, so that the temperature of the internal device 14 can be appropriately maintained during the flight of the two-stage rocket 1 and the orbital operation of the artificial satellite 2. It is possible.
[0051]
In the present embodiment, the diameter of the ventilation hole 15 is 1 mm, and the distance between two adjacent ventilation holes 15 is 25 mm. In addition, the diameter of the ventilation hole 15 is set to 3 mm, and the distance between two adjacent ventilation holes 15 is set to 5 mm. The ventilation hole 15 reduces the exposed area of the heat protection layer 8a as much as possible. It goes without saying that any other dimensional configuration may be used as long as the configuration has the desired performance of releasing the pyrolysis gas to the outside.
[0052]
Next, an example of a method for manufacturing the fairing 6 according to the embodiment of the present invention will be described. 7A and 7B are schematic views illustrating an example of a pre-process of painting the fairing 6, wherein FIG. 7A illustrates the first stage, FIG. 7B illustrates the second stage, and FIG. 7C illustrates the third stage. And (d) shows the fourth stage. First, a fairing material 16 which is a laminate of the ablator layer 8, the heat insulating layer 9, and the metal sheet layer 10 molded into the shape of the fairing 6 is manufactured. Then, a paint having a low solar absorptivity (for example, white) is applied to the surface of the fairing material 16 as follows.
[0053]
As shown in FIG. 7A, in the first step of the pre-process of painting the fairing 6 in this example, the adhesive sealing material 17 that has been manufactured in advance according to the outer shape of the fairing 6 is similarly removed. It is put on a female jig 18 that has been manufactured according to the outer shape of the fairing 6 in advance. The female jig 18 is provided with a large number of holes 18a in advance corresponding to the positions of the ventilation holes 15 provided in the low solar absorptivity layer 7 of the fairing 6.
[0054]
Next, in the second stage, as shown in FIG. 7B, the female jig 18 is put on the fairing material 16.
[0055]
In the third stage, as shown in FIG. 7C, a male jig 19 that has been manufactured in advance according to the outer shape of the fairing 6 is put on the sealing material 17. On the inner surface of the male jig 19, a number of protrusions 19a are provided in advance in accordance with the positions of the ventilation holes 15 provided in the low solar absorptivity layer 7 of the fairing 6. Then, the positioning is performed so that the projection 19a is inserted into the hole 18a, and the male jig 19 is pressed against the fairing material 16. At this time, the portion of the sealing material 17 overlapping the projection 19a penetrates the hole 18a and comes into contact with the surface of the fairing material 16, and the sealing material 17 is adhered only to a position corresponding to the ventilation hole 15 of the fairing material 16. Can be
[0056]
In the fourth stage, as shown in FIG. 7D, the male jig 19, the sealing material 17 and the female jig 18 are removed from the fairing material 16. At this time, only a portion of the seal material 17 where the fairing material 16 is adhered remains on the surface of the fairing material 16 as a masking (masking portion) 17a, and this portion of the ablator layer 8 is hidden. Thus, the pre-process of painting the fairing 6 is completed.
[0057]
A heat-resistant paint having a low solar absorptivity is applied to the outer surface of the fairing material 16 thus obtained. Thereafter, the masking 17a is removed from the fairing material 16, and the fairing 6 is completed.
[0058]
Next, another example of the method for manufacturing the fairing 6 according to the embodiment of the present invention will be described. FIGS. 8A and 8B are schematic diagrams illustrating another example of the pre-process of the painting of the fairing 6, in which FIG. 8A shows the first stage and FIG. 8B shows the second stage.
[0059]
As shown in FIG. 8A, in the first stage of the pre-process of painting the fairing 6 in this example, the electromagnet 20 that has been manufactured in advance according to the inner shape of the fairing 6 is placed inside the fairing material 16. Deploy.
[0060]
FIG. 9 is a partial cross-sectional view showing the configuration of the electromagnet 20 according to the embodiment of the present invention. The electromagnet 20 includes a body portion 21 in which a plurality of coils (not shown) are provided, and has a substantially similar shape obtained by reducing the fairing material 16, and is disposed outside and inside the body portion 21, respectively. And the honeycomb structure portions 22 and 23. The outer shape of the outer honeycomb structure portion 22 is substantially the same as the inner shape of the fairing material 16, and when the electromagnet 20 is arranged inside the fairing material 16, the honeycomb structure portion 22 is formed on the inner surface of the fairing material 16. Will be in close contact over substantially the entire surface.
[0061]
The honeycomb structure portions 22 and 23 are made of metal, and have a structure in which a plurality of hexagonal holes penetrate in the thickness direction and are formed in a honeycomb shape. With the honeycomb structure portions 22 and 23, the electromagnet 20 can be made lightweight and highly rigid.
[0062]
Further, the coil provided on the main body portion 21 has the center axis in the thickness direction of the main body portion 21, and when a direct current is supplied to the coil, for example, as shown in FIG. Can be generated as an N pole and an inner S pole.
[0063]
In the present embodiment, the electromagnet 20 is arranged inside the fairing material 16. However, the present invention is not limited to this, and a permanent magnet may be arranged instead of the electromagnet 20. Further, the electromagnet 20 is configured to have the honeycomb structure portions 22 and 23, but is not limited to this, and is made of a metal or a metal and a synthetic resin having the same outer shape as the inner shape of the fairing material 16. A configuration in which a plurality of coils are housed inside a composite housing may be employed.
[0064]
Next, in the second stage of the present example, as shown in FIG. 8B, a paint supply device 24 whose inside is manufactured in advance to substantially match the outer shape of the fairing 6 is put on the outside of the fairing material 16.
[0065]
FIG. 10 is a partial perspective view showing the configuration of the paint supply device 24 according to the embodiment of the present invention. The paint supply device 24 includes a bag-shaped main body portion 25 made of a flexible material, which is larger than the fairing material 16, and a large number of paints protruding from a plurality of locations on the inner surface of the main body portion 25. It is mainly composed of an injection tube 26 and a column 27 protruding from a plurality of locations different from the projection location of the paint injection tube 26 on the inner surface of the main body portion 25.
[0066]
The main body portion 25 has an internal space 28 through which heat-resistant paint having a low solar absorptivity and supplied from the outside can flow. As described above, the main body portion 25 has flexibility, and is formed in a shape substantially conforming to the outer shape of the fairing material 16. Therefore, when the fairing material 16 is covered by the paint supply device 24, the main body portion 25 is deformed according to the outer shape of the fairing material 16, and covers the main body portion 25 over substantially the entire fairing material 16. Is possible.
[0067]
FIG. 11 is a schematic side view showing the configuration of the paint spray tube 26 according to the embodiment of the present invention. One end of the paint spray tube 26 communicates with the internal space 28, and the other end is closed by a masking portion 29 made of a substantially disk-shaped permanent magnet. Further, a plurality of nozzles 30 are provided in an intermediate portion of the paint spray tube 26 so that the heat-resistant paint supplied from the internal space 28 can be sprayed from the nozzles 30. At this time, since the other end of the paint spray tube 26 is closed by the masking portion 29, the heat-resistant paint is not ejected from this other end.
[0068]
A support 27 made of a material having high magnetic permeability projects from the inner surface of the main body 25. One end of the column 27 is fixed to the inner surface of the main body 25, and a masking portion 32 is attached to the other end. This masking section 32 is also a permanent magnet. For example, when the electromagnet 20 generates a magnetic field in which the outside of the main body portion 21 is an N pole and the inside is an S pole, the masking portions 29 and 32 are an N pole near the main body portion 25 and an S pole on the far side. They are attached to the paint spray tube 26 and the column 27, respectively, so as to generate a poled magnetic field.
[0069]
When the paint supply device 24 is put on the fairing material 16 in the second stage, the masking portions 29 and 32 are placed on the outer surface of the fairing material 16 where the air holes 15 are to be provided. Then, a magnetic field is generated by the electromagnet 20, and the masking portions 29 and 32 are attracted to the outer surface of the fairing material 16 by the magnetic field. Thus, the pre-process of painting the fairing 6 is completed.
[0070]
Subsequently, the heat-resistant paint is supplied from the outside of the paint supply device 24 to the main body portion 25, and the heat-resistant paint is jetted from the nozzle 30. The nozzle 30 is provided so as to be inclined so as to spray the heat-resistant paint toward the surface of the fairing material 16, whereby the masking portions 29 and 32 on the outer surface of the fairing material 16 are placed. The heat-resistant paint is applied over substantially the entire area except for the above.
[0071]
Further, the generation of the magnetic field of the electromagnet 20 is stopped, and the electromagnet 20 and the paint supply device 24 are removed from the fairing material 16. Thus, the fairing 6 is completed.
[0072]
As described above, by configuring the artificial satellite 2 in which the fairing 6 is integrated with the outer shell, a fairing is prepared separately from the artificial satellite like a conventional artificial satellite, and the artificial satellite is provided together with the fairing. There is no need to mount it on a rocket. An artificial satellite requires an outer shell structure, and in order to secure sufficient strength and rigidity of the entire artificial satellite, a conventional artificial satellite also requires a strength member and the like. In the artificial satellite 2 according to the present embodiment, the fairing 6 also serves as an outer shell structure, and since the fairing 6 has sufficient strength and rigidity as described above, it is necessary to separately provide a strength member or the like. Absent.
[0073]
Further, as a result of trial calculations by the inventors of the present application, the fairing 6 according to the embodiment of the present invention mainly constituted by CFRP having a smaller specific gravity than metal has a conventional artificial satellite equipped with the same internal equipment. Compared to a metal fairing used when launching a rocket with a three-stage rocket, it is found that when both are configured to secure the same strength and rigidity, they have the same mass. Was. That is, the artificial satellite 2 according to the present invention requires the conventional artificial satellite when the internal equipment mounted on the artificial satellite is the same as the total mass of the conventional artificial satellite and the metal fairing. It is possible to reduce the weight by the outer shell structure and the mass of the strength member. For example, when the mass of the conventional artificial satellite is 50 kg, the mass of the outer shell structure and the strength member required by the artificial satellite is about 15 kg, and therefore, the artificial satellite 2 according to the present embodiment equipped with the same internal device is mounted. In this case, the weight can be reduced by 15 kg from the total mass of the conventional artificial satellite and the fairing.
[0074]
By reducing the weight of such satellites (and fairings), the rocket itself that transports the satellites to outer space can be made smaller. In addition, as in the present embodiment, instead of the conventional multi-stage rocket having three or more stages, the two-stage rocket 1 is used, so that the structure can be greatly simplified. Etc. can be greatly reduced.
[0075]
As described above, when the fairing 6 is integrated with the artificial satellite 2, the conventional artificial satellite and the fairing are separately mounted, and the same rocket as that configured to separate and throw the fairing is used. It is possible to launch the artificial satellite 2 having a mass about 15 kg larger than the artificial satellite using a rocket. Paradoxically, a fairing-integrated satellite having the same mass as a conventional artificial satellite configured to separate and release a fairing can be launched with a significantly smaller rocket than before.
[0076]
Further, as described above, the two-stage rocket 1 flies at a much higher speed than a conventional multi-stage rocket having three or more stages. For example, when the gravitational turn of about 2 ° / s is started 39 seconds after the start of combustion of the first-stage solid motor 3 and the angle of attack is 90 ° at the end of combustion of the second-stage solid motor 4, The horizontal movement distance of the two-stage rocket 1 at the end of the combustion of the two-stage solid motor 4 is about 180 km, and when the artificial satellite 2 is to be released 10 seconds later, the two-stage rocket 1 (the artificial satellite 2) ) Is about 260 km.
[0077]
As described above, the horizontal movement distance from the start of launch of the artificial satellite 2 to the time of launch is about 260 km, which is extremely shorter than the horizontal travel distance of the conventional artificial satellite 2 from the launch to launch of the conventional satellite. . Therefore, when a conventional three-stage rocket launches an artificial satellite, for example, after launching from a launch site in Japan, it may not be possible to observe from launch to release of the satellite by an observatory in Japan. When the two-stage rocket 1 equipped with the artificial satellite 2 according to the present invention is launched from a small rocket launch site at the launch site of the Tanegashima Space Center, the flight status of the two-stage rocket 1 until the release of the artificial satellite 2 will be described. Observations can be made at two locations, Ogasawara Tracking Station, and no foreign facilities are required for observation. As described above, since the observation from launching of the two-stage rocket 1 to orbiting of the satellite 2 can be performed only with the facilities in Japan, the launching work of the rocket is extremely simplified as compared with the conventional case. be able to.
[0078]
The horizontal movement distance (about 260 km) from the start of the launch of the artificial satellite 2 according to the present invention to the release thereof is measured from the launch of the two-stage rocket 1 to the insertion of the artificial satellite 2 into orbit by one observatory. It is also the distance you can do. Therefore, when the observation from the launching of the two-stage rocket 1 to the orbiting of the artificial satellite 2 is performed consistently at one observatory, the launching work of the rocket can be further simplified, and the cost can be further reduced. In addition, it can be expected to reduce labor.
[0079]
As described above, in the two-stage rocket 1 according to Embodiment 2, the ablator layer 8 of the fairing 6 is carbonized by aerodynamic heating. At this time, the carbon fibers remain in a portion near the surface of the ablator layer 8 in an overlapping state. FIG. 12 is a schematic cross-sectional view illustrating a state when ablator layer 8 according to the embodiment of the present invention has been subjected to aerodynamic heating. The heat protection layer 8a of the ablator layer 8 is carbonized at the initial stage of the launch of the two-stage rocket 1. As shown in FIG. 12, the carbon fibers 8c carbonized and remaining on the heat protection layer 8a overlap with the fiber longitudinal directions oriented in various directions. The bonding force of each carbon fiber 8c is relatively weak, and when a force is applied from the outside, the carbon fibers 8c move slightly so as to be shifted from each other.
[0080]
Here, as described above, the two-stage rocket 1 according to the present embodiment receives an extremely high dynamic pressure at the time of its flight as compared with the related art. Accordingly, strong vibration is generated in the two-stage rocket 1 during flight, and if the vibration applied to the fairing 6 is transmitted to the internal device 14 as it is, this may cause a failure of the internal device 14.
[0081]
However, the carbon fibers 8b (and the carbon fibers remaining in this layer when the structural layer 8c is carbonized) move three-dimensionally and minutely, so that the ablator layer 8 functions as a seismic isolation structure, 6 is prevented from being transmitted to the inside as it is. This makes it difficult for external vibrations to be transmitted to the internal device 14, and it is possible to suppress occurrence of a failure in the internal device 14.
[0082]
When the artificial satellite 2 is put into orbit, the second-stage solid-state motor 4 and the inter-stage portion 5 are separated from the artificial satellite 2, but an impact is applied to the artificial satellite 2 also at this time. This impact is also mitigated inside the artificial satellite 2 by the seismic isolation function of the fairing 6 described above, and the occurrence of failure of the internal device 14 is suppressed.
[0083]
Next, the configuration of the artificial satellite 2 will be described in more detail. FIG. 13 is a cross-sectional view of the shroud portion 2a showing the configuration of the artificial satellite 2 according to the embodiment of the present invention. As shown in FIG. 13, in a portion from the rear end of the shroud portion 2 a of the artificial satellite 2 to the front of a predetermined length, a part of the fairing 6 in the circumferential direction is missing, and the missing portion 33 is closed. Is provided with a deployable solar cell paddle 34.
[0084]
The solar battery paddle 34 has a configuration in which two plate-like solar battery panels 35 and 36 are abutted and hinged to each other. One end of one solar cell panel 35 is bent in a substantially L shape, and the tip of the bent portion 34 b is pivotally connected to one end of the missing portion 33. This allows the solar cell panel 35 to pivot outward around the pivoted portion from a state in which the missing portion 33 is closed.
[0085]
The other end of the solar cell panel 35 is hinged to one end of the solar cell panel 36 by a hinge 34a. Before the launch of the two-stage rocket 1, the solar cell panels 35 and 36 are housed in a superposed state so as to close the missing portion 33.
[0086]
FIG. 14 is a cross-sectional view showing a configuration of solar cell paddle 34 according to the embodiment of the present invention. As shown in FIG. 14, one (base end) solar cell panel 35 has a structure in which a solar cell 35b is mounted on one surface of a cell mounting plate 35a. A radiator 35c is provided so as to close the missing portion 33, and the radiator 35c functions as a partition between the inside of the artificial satellite 2 and the outer space. Discharge of the heat generated by the internal device 14 to the outside is performed by the radiator 35c. The optical characteristics and area of the radiator 35c are set based on the amount of heat generated by the internal device 14. When the solar battery paddle 34 is stored, the other surface of the cell mounting plate 35a faces the outer surface of the radiator 35c, and the solar battery cells 35b are arranged on the outer device.
[0087]
In the present embodiment, the configuration in which the radiator 35c is provided as a partition between the inside of the artificial satellite 2 and the outer space has been described. However, the present invention is not limited to this, and the amount of heat generated by the internal device 14 depends on the radiator. When the exhaust heat is unnecessary and is small to some extent, a configuration may be adopted in which a partition made of a composite material such as CFRP is provided instead of the radiator 35c.
[0088]
The other (front end) solar cell panel 36 has a cell mounting plate 36a, and when the solar cell paddle 34 is stored, the solar cell 36b faces the solar cell 35b. 36a is attached to one surface. That is, the solar cell 36b is mounted on the inner surface of the cell mounting plate 36a. Further, a heat protection plate 36c composed of an ablator layer, a heat insulating layer, and the like is mounted on the other surface (surface on the outside of the machine) of the cell mounting plate 36a. This makes it possible to exhibit a thermal protection function against aerodynamic heating received when the two-stage rocket 1 flies, so that the temperature of the internal device 14 is not excessively increased. When the solar battery paddle 34 is stored, the thermal protection plate 36c is located at the outermost position, so that the solar battery cells 35b and 36b are protected from aerodynamic heating received when the two-stage rocket 1 flies.
[0089]
As shown in FIG. 13, two electric motors 37 and 38 are provided inside the artificial satellite 2. One end of a wire 37a is connected to the output shaft of one electric motor 37, and the other end of the wire 37a is connected to a pivot 35d that pivotally connects the solar cell panel 35 and the fairing 6. One end of the wire 37a is wound around the output shaft of the electric motor 37, and the other end is wound around the pivot 35d in the opposite direction. Further, the pivot 35d is fixed to the solar cell panel 35, and the winding direction of the other end of the wire 37a around the pivot 35d is such that when the wire 37a is pulled out from the pivot 35d, the solar cell panel 35 is turned outward. The direction in which the torque in the moving direction is generated (that is, the counterclockwise direction in FIG. 13). Thereby, when the electric motor 37 is operated and the output shaft is rotated in the direction of winding the wire 37a, the wire 37a is pulled out from the pivot 35d, and the solar cell panel 35 is turned in the direction of turning outside. The torque acts.
[0090]
On the other hand, one end of the wire 38a is also wound around the output shaft of the electric motor 38. The other end of the wire 38a is connected to the hinge 34a. When the solar battery paddle 34 is stored in the fairing 6, the wire 38a is fixed by the output shaft of the electric motor 38 so that a predetermined tension acts. Further, the solar cell paddle 34 may be locked to the fairing 6 by a lock mechanism so as not to be expanded, and the lock mechanism may be released when the artificial satellite 2 is orbited.
[0091]
Further, inside the artificial satellite 2, there are provided a motor controller 39 for controlling operations of the electric motors 37 and 38, respectively, and an acceleration sensor 40 connected to the motor controller 39.
[0092]
A spring (not shown) is provided between the solar cell panels 35 and 36, and the spring urges the solar cell panels 35 and 36 in a direction in which the solar cell panels 35 and 36 are developed. Further, a stopper (not shown) is provided between the solar cell panels 35 and 36. When the solar cell paddle 34 is stored in the fairing 6, the stopper engages with the solar cell panels 35 and Numeral 36 maintains the state of being superimposed against the urging force of the spring.
[0093]
Next, the deployment operation of the solar battery paddle 34 will be described. FIG. 15 and FIG. 16 are schematic diagrams for explaining a flow of deployment of the solar cell paddle 34 according to the embodiment of the present invention. When the combustion of the second-stage solid motor 4 is completed, the acceleration sensor 40 detects this and transmits a detection signal to the motor controller 39. At the same time, the second-stage solid motor 4 and the interstage 5 are disconnected from the artificial satellite 2, and the artificial satellite 2 is released into orbit. After receiving the detection signal, the motor controller 39 operates the electric motor 37 so that its output shaft rotates in the winding direction of the wire 37a, and controls the electric motor so that the output shaft rotates in the direction of loosening the wire 38a. The motor 38 is operated. As a result, the solar cell paddle 34 rotates outward about the pivot 35d. When the solar battery paddle 34 is rotated, the missing portion 33 is closed by the radiator 35c only.
[0094]
As shown in FIG. 15, the motor controller 39 drives the electric motors 37 and 38 until the solar battery paddle 34 has just rotated by 90 °. When the solar cell paddle 34 rotates 90 ° from the start of rotation, the engagement of the stopper is released, and the solar cell panel 36 is centered on the hinge portion 34a by the biasing force of the spring described above. As a result, it turns in the direction away from the solar cell panel 35 (see FIG. 16). The rotation of the solar cell panel 36 further proceeds due to the inertial force, and when the solar cell panel 36 is rotated by 180 ° to be in a state where the solar cell panels 35 and 36 are linearly arranged, the solar cell panels 35 and 36 are arranged. The stoppers (not shown) provided in the vicinity of the above are engaged so that the solar cell panels 35 and 36 are maintained in a state of being linearly arranged. Thus, the operation of deploying the solar battery paddle 34 is completed.
[0095]
As described above, since the electric motors 37.38 and the solar battery paddles 34 are connected by the lightweight wires 37a and 38a, and the electric motors 37 and 38 are driven, the solar battery paddles 34 are deployed. The artificial satellite 2 can be significantly reduced in weight as compared with a configuration in which the rotational motion of the output shafts of the electric motors (actuators) 37 and 38 is transmitted to the solar battery paddle 34 using such a mechanism.
[0096]
FIG. 17 is a perspective view showing a state during orbital operation of artificial satellite 2 according to the embodiment of the present invention. As shown in FIG. 17, a solar sensor 41 is provided near the tip of the solar cell 36b of the solar cell paddle 34. The sun sensor 41 outputs an electric signal corresponding to the solar radiation intensity, and the output destination is the motor controller 39. The structure of the sun sensor 41 will be described in more detail. A so-called coarse sensor (CSSA) is used as the sun sensor 41. The sensor allows sunlight to enter through a slit and the incident angle of the sunlight. This is a configuration in which only the component in the direction orthogonal to the slit is measured, and is known. The sun sensor 41 is arranged such that the light receiving surface of the solar battery paddle 34 and the light receiving surface thereof are parallel. The motor controller 39 drives the electric motors 37 and 38 so that the light receiving surface of the sun sensor 41 faces the sun, and thereby the solar cell paddle 34 causes the light receiving surface of the solar cell paddle 34 to face the sun. Is controlled. At this time, it is desirable that the light receiving surface of the solar cell paddle 34 is completely opposed to the sun. However, due to the relationship between the orbital plane and the sunlight vector, it is rare that the light receiving surface is completely opposed to the sun. The deployment angle of the solar cell paddle 34 is controlled so as to face the same. Then, the solar cells 35b and 36b receive the solar radiation and the earth albedo (earth reflection component of the solar radiation) to generate power, and the power obtained by the solar cells 35b and 36b is supplied to the internal device 14. It has become.
[0097]
Assuming that the mass of the artificial satellite 2 is 45 kg, the required power can be estimated to be approximately 20 W from the past statistical results on the relationship between the mass of the artificial satellite and the power used by the internal devices. Since the solar cells 35b and 36b cannot generate power in the night orbit, if the day orbit time and the night orbit time are the same, the required power generation is about 40 W, which is twice the required power. . There are two types of photovoltaic cells, which are widely used in space, a gallium arsenide type, which is expensive but has high efficiency, and a silicon type, which is inexpensive but has relatively low power generation efficiency. If silicon-type solar cells are used as the solar cells 35b and 36b, the light-receiving surface area required for power generation of 80 W which is a value obtained by multiplying the required power generation by 2 as a safety factor is about 0.6 m. ^Two It becomes. As a result of the calculation by the present inventors, the area of the solar battery paddle 34 of the artificial satellite 2 having a mass of 45 kg is 0.9 m. ^Two Therefore, a required amount of power can be obtained even when a silicon-type solar cell is used.
[0098]
When the artificial satellite 2 is put into orbit, the S-band antenna or the gravitational tilt beam housed inside the artificial satellite 2 is deployed (not shown) together with the deployment of the solar battery paddle 34 if necessary. The orbital injection of the satellite 2 is completed.
[0099]
As shown in FIG. 17, the artificial satellite 2 travels in the direction of the tip of the nose portion 2b, but a small amount of air exists in the space around the artificial satellite 2, and the artificial satellite 2 has an air resistance. Will be received. When the main body of the satellite has a hexahedron or cylindrical shape as in the case of a conventional satellite, the effect of such air resistance is large, the satellite is decelerated, and the balance between centrifugal force and gravity is reduced. The satellite will collapse and the altitude of the satellite will decrease. On the other hand, in the artificial satellite 2 according to the present invention, since the fairing 6 which is the main body of the artificial satellite has a streamline shape, it is less likely to receive air resistance than the conventional artificial satellite, and the air resistance of the artificial satellite 2 is small. The deceleration due to is suppressed as much as possible.
[0100]
However, the artificial satellite 2 according to the present invention continues to receive some air resistance and decelerates, and when the altitude decreases, it is necessary to increase the thrust by the thrust generated by the thrust generator. Conventional large-sized, small-sized, or mini-type artificial satellites use a thrust generator configured to burn liquid fuel or to inject liquid nitrogen or the like. If the target operating life of the satellite is set to one year or more, a large amount of fuel or injection material is required to maintain the altitude of the satellite. In addition, such a thrust generator has a complicated configuration, and the size and mass of the device itself are large. Therefore, a large, small, or mini type satellite has a relatively large equipment mounting volume, and it is relatively easy to mount such a conventional thrust generator. Micro or pico satellites have a small equipment mounting volume, and it is not easy to mount these thrust generators on micro or pico satellites.
[0101]
Therefore, the artificial satellite according to the present embodiment is equipped with a thrust generator as described below. FIG. 18 is a schematic sectional view showing the configuration of the thrust generating device according to the present embodiment. As shown in FIG. 18, the thrust generator 42 according to the present embodiment is mainly configured by a case 43, a pipe 44, a nozzle 45, a heater 46, an on-off valve 47, temperature sensors 48a and 48b, and a pressure sensor 49. The case 43 has a rectangular parallelepiped box shape, in which a phase change substance 50 is stored. As the phase change material 50, eicosan or the like having a low melting point of about 300K is used.
[0102]
One surface of the case 43 is open, and a pipe 44 extends from the opening. The end of the pipe 44 is connected to a vertex portion of a nozzle 45 whose diameter has been expanded to a substantially conical shape. An on-off valve 47 is provided in the middle of the pipe 44, and when the on-off valve 47 is closed, a portion from the connection point of the case 43 and the pipe 44 to the case 43 to the on-off valve 47. Constitutes a closed space, and when the on-off valve 47 is open, the space in the case 43 communicates with the outside through the pipe 44 and the nozzle 45.
[0103]
Further, a heater 46 is provided so as to cover substantially the entire outer peripheral surfaces of the case 43, the pipe 44, and the nozzle 45. As the heater 46, a type such as a heating wire that converts electric power into heat is used, and the case 43, the pipe 44, and the nozzle 45 are heated by receiving electric power from the outside.
[0104]
A temperature sensor 48 a and a pressure sensor 49 are provided inside the case 43, and a temperature sensor 48 b is also provided in a middle portion between the nozzle 45 and the on-off valve 47 of the pipe 44.
[0105]
Next, the operation of the thrust generating device 42 according to the present embodiment will be described. When it is determined by the on-board computer (not shown) built in the artificial satellite 2 that the artificial satellite 2 needs thrust, electric power is supplied to the heater 46 from the outside, and the case 43, the pipe 44 and The nozzle 45 is heated. The temperature of the internal space of the case 43 rises, and the sublimation gas 51 of the phase change substance 50 is generated (increased), and the internal pressure of the case 43 increases. An electric signal corresponding to the internal pressure of the case 43 is transmitted to the on-board computer by the pressure sensor 49, and the on-board computer stops supplying power to the heater 46 when the internal pressure of the case 43 becomes a predetermined value or more. At the same time, the operation is controlled to open the on-off valve 47. When the on-off valve 47 is opened, the sublimation gas 51 in the case 43 is injected from the nozzle 45, and a thrust in the direction opposite to the injection direction of the sublimation gas 51 is generated.
[0106]
After the injection of the sublimation gas 51, the on-board computer monitors the output values of the temperature sensors 48a and 48b, and the case 43, the pipe 44, the nozzle 45, and the on-off valve 47 are reduced to a predetermined low-temperature side allowable temperature or less by adiabatic expansion of the gas. If necessary, these are heated by the heater 46 so as not to be rapidly cooled. After the injection of the sublimation gas 51, the on-board computer closes the on-off valve 47, monitors the output value of the temperature sensor 48a, and controls the temperature of the phase change substance 50 in an appropriate range. For example, the case 43 is heated by the heater 46.
[0107]
Next, the result of trial calculation of the required amount of the phase change substance 50 will be described. In the following description, the cross-sectional area viewed from the traveling direction is 0.7 m ^Two The result of trial calculation is shown for a case where the altitude of the artificial satellite 2 is 350 km, where the air resistance is relatively large. According to the calculation by the inventor of the present application, when the mass of the artificial satellite is 45 kg, the amount of deceleration per day can be canceled by the ejection of the sublimation gas 51 of 15 g. In this case, if the operational life of the artificial satellite 2 is 100 days, the required mass of the phase change substance 50 is 1.5 kg. For example, if eicosane is used for the phase change substance 50, the density of eicosane is 830 kg / m ^Three Therefore, the required volume of the case 43 is 1.8 × 10 ^-3 m ^Three This is a size that can be sufficiently mounted on the artificial satellite 2.
[0108]
Further, the thrust generated by the thrust generator 42 according to the present embodiment is significantly smaller than that of the conventional thrust generator, but since the mass of the artificial satellite 2 is at most several tens of kilograms, The purpose used for maintaining the altitude and speed of the artificial satellite 2 can be sufficiently achieved.
[0109]
Next, a configuration relating to attitude control during orbital operation of the artificial satellite 2 according to the present embodiment will be described. FIG. 19 is a schematic perspective view showing an example of the configuration of the attitude control mechanism in the pitching direction of the artificial satellite 2 according to the embodiment of the present invention. At the front end (the downstream end in the traveling direction of the artificial satellite 2) of the solar cell panels 35 and 36, square tubular guide tubes 52a and 52b are provided along the front edges of the solar cell panels 35 and 36, respectively. Similarly, guide tubes 53a and 53b are provided at the rear ends of the solar panels 35 and 36 (upstream ends in the traveling direction of the artificial satellite 2) along the rear edges of the solar panels 35 and 36, respectively. . As shown in FIG. 19, the guide tubes 52a and 52b are configured so as to be linearly arranged when the solar battery paddle 34 is deployed, and similarly, the guide tubes 53a and 53b are The paddles 34 are arranged linearly when the paddles 34 are deployed. A wire 52c is inserted into the guide tubes 52a and 52b arranged in a straight line, and the inside end of the wire 52c is attached to an output shaft of an electric motor 54 provided inside the satellite 2. I have. A weight 52d is attached to the outside end of the wire 52c.
[0110]
Similarly, a wire 53c is inserted into the guide tubes 53a and 53b arranged in a straight line, and the inside end of the wire 53c is connected to the output shaft of an electric motor 55 provided inside the artificial satellite 2. Installed. A weight 53d is attached to the outside end of the wire 53c.
[0111]
The electric motors 54 and 55 are connected to a motor controller (not shown) built in the artificial satellite 2, and are configured to be operated and controlled by the motor controllers. As shown in FIG. 17, an earth sensor 56 is provided so as to project from the artificial satellite 2 toward the earth, and the motor controller is connected to the earth sensor 56.
[0112]
The earth sensor 56 outputs an electric signal corresponding to each of the roll angle and the pitch angle of the artificial satellite 2 with respect to the earth. Among them, the output signal representing the pitch angle is received by the motor controller, and the motor controller receives the output signal. The operation of the electric motors 54 and 55 is controlled in order to adjust the length of the wires 52c and 53c. More specifically, for example, proportional control is performed so that the pitch angle of the satellite detected by the earth sensor 56 approaches a predetermined target pitch angle (that is, a pitch angle of 0 ° in the present embodiment). The rotation direction and rotation speed of the electric motors 54 and 55 are controlled using a known control method such as PI control, PID control, or fuzzy control. Since the weights 52d and 53d receive the gravity of the earth, the wires 52c and 53c respectively hang down to the earth. When the operation of the electric motors 54 and 55 is controlled, when the wire 52c is pulled out longer than the wire 53c, the weight 52d comes closer to the earth than the weight 53d. As a result, the weight 52d receives the gravity from the earth more strongly than the weight 53d. As a result, the position of the center of gravity of the artificial satellite 2 moves forward, the leading end side of the artificial satellite 2 approaches the earth, and the rear end side moves. Attitude control of the artificial satellite 2 in the pitching direction is performed so as to separate from the earth. On the other hand, when the operation of the electric motors 54 and 55 is controlled and the wire 53c is pulled out longer than the wire 52c, the position of the center of gravity of the artificial satellite 2 moves backward, and the leading end side of the artificial satellite moves away from the earth, The attitude control of the artificial satellite 2 in the pitching direction is performed so that the rear end side approaches the earth.
[0113]
The operation of the electric motors 54 and 55 is controlled so as to maintain the attitude of the artificial satellite 2 in the pitching direction so that the earth sensor 56 always faces the center of the earth.
[0114]
Further, the guide tubes 52a, 52b, 53a, 53b can also be used as strength members of the solar battery paddle 34.
[0115]
FIG. 20 is a schematic perspective view showing another example of the configuration of the attitude control mechanism in the pitching direction of the artificial satellite 2 according to the embodiment of the present invention. In the present example, as shown in FIG. 20, a moving weight 57 is provided on the outer edge of the solar cell paddle 34 outside the artificial satellite 2, that is, on the opposite edge of the hinge portion 34a of the solar cell panel 36. I have. The moving weight 57 is attached to the runner 58b of the linear motor 58 having a stator 58a and a runner 58b provided over the entire length of the edge of the solar cell panel 36. Along with the traveling element 58b.
[0116]
The linear motor 58 is connected to a motor controller (not shown), and the motor controller controls the operation of the linear motor 58 in accordance with an output signal from the earth sensor 56 relating to the pitch angle of the artificial satellite 2. ing. Here, the operation of the linear motor 58 is controlled by the known control method as described above. When the moving weight 57 moves in the front-rear direction of the artificial satellite 2, the position of the center of gravity of the artificial satellite 2 moves, and thereby the attitude control of the artificial satellite 2 in the pitching direction is performed.
[0117]
In addition, it is good also as a structure provided with the movable weight which can move to the front-back direction inside the fairing 6. FIG.
[0118]
The attitude control of the satellite 2 in the rolling direction is performed by adjusting the deployment angle of the solar battery paddle 34 by the motor controller 39 controlling the operation of the electric motors 37 and 38. Specifically, the motor controller 39 receives an output signal indicating the roll angle of the artificial satellite 2 from the earth sensor 56, and the motor controller 39 converts the detected roll angle of the artificial satellite 2 into a predetermined target roll angle. In the case of the embodiment, the rotation directions and the rotation speeds of the electric motors 37 and 38 are adjusted using a known control method such as proportional control, PI control, PID control, or fuzzy control so as to approach 0 ° roll angle. Control. At this time, the solar battery paddle 34 is tilted to the opposite side of the direction in which the satellite 2 is tilted with respect to the center of the earth, thereby moving the position of the center of gravity of the satellite 2 in the horizontal direction, and consequently the roll angle is reduced to 0 ° The attitude control of the artificial satellite 2 in the rolling direction is performed by inclining the artificial satellite 2 in a direction approaching.
[0119]
Further, during the orbital operation of the artificial satellite 2, the motor controller 39 controls the operation of the electric motors 37 and 38 so that the deployment angle of the solar battery paddle 34 is adjusted so that the light receiving surface of the solar battery paddle 34 faces the sun. Controlled. Thereby, the solar cell paddle 34 can be operated alone and its light receiving surface can be directed to the sun, and it is necessary to control the attitude of the satellite itself in order to direct the light receiving surface of the solar cell paddle to the sun as in the related art. There is no.
[0120]
【The invention's effect】
In the case of the artificial satellite according to the present invention, the cross-sectional area at the tip of the artificial satellite in the traveling direction is significantly reduced as compared with the conventional one, and the air resistance received by the artificial satellite in orbit is reduced as much as possible. Can be.
[0121]
Further, since the outer shell of the artificial satellite is used as a fairing, it is not necessary to provide a fairing separately from the artificial satellite, and it is possible to launch using a rocket having a size optimal for the artificial satellite. Therefore, in the artificial satellite according to the present invention, there is no extra space between the artificial satellite and the fairing, and the spatial efficiency at the time of mounting the artificial satellite on the rocket used for the launch is significantly higher than the conventional one. As a result, drastic miniaturization and cost reduction of the entire rocket for launching a satellite can be realized.
[0122]
Further, in the artificial satellite according to the present invention, since the outer shell of the artificial satellite and the fairing are also used, the mass of the artificial satellite is significantly reduced as compared with the total mass of the conventional artificial satellite and the fairing. The weight can be reduced. This makes it possible to significantly reduce the size and weight of the launch rocket as compared to conventional satellites.
[0123]
In addition, by configuring such a small and lightweight launch vehicle, the amount of launch fuel of the launch vehicle can be significantly reduced.
[0124]
Further, since the launch rocket can be configured to be small and light, the costs involved in the development, manufacture, and launch of the rocket can be significantly reduced as compared with the conventional rocket, and thus the present invention can be applied. The present invention has an excellent effect, for example, it can be used exclusively for launching a configured pico-type or micro-type artificial satellite.
[Brief description of the drawings]
FIG. 1 is a side view showing the overall configuration of a two-stage rocket equipped with an artificial satellite according to an embodiment of the present invention.
FIG. 2 is a schematic cross-sectional view schematically illustrating a configuration of an artificial satellite according to an embodiment of the present invention.
FIG. 3 is a cross-sectional view illustrating a configuration of an ablator layer according to the embodiment of the present invention.
FIG. 4 is a graph showing changes in the temperature of the surface of the fairing, the temperature of the metal sheet layer, and the temperature of the internal equipment when the two-stage rocket according to the embodiment of the present invention flies.
FIG. 5 is a partial perspective view of a fairing showing a configuration of a low solar absorptivity layer according to the embodiment of the present invention.
FIG. 6 is a schematic partial sectional view of a fairing according to the embodiment of the present invention.
FIGS. 7A and 7B are schematic diagrams illustrating an example of a pre-process of painting a fairing according to an embodiment of the present invention, wherein FIG. 7A illustrates a first stage, FIG. (c) is a schematic diagram showing the third stage, and (d) is a schematic diagram showing the fourth stage.
FIGS. 8A and 8B are schematic diagrams illustrating another example of a pre-process of painting a fairing according to the embodiment of the present invention, wherein FIG. 8A illustrates a first stage, and FIG. It is a schematic diagram shown respectively.
FIG. 9 is a partial cross-sectional view illustrating a configuration of an electromagnet according to an embodiment of the present invention.
FIG. 10 is a partial perspective view illustrating a configuration of a paint supply device according to the embodiment of the present invention.
FIG. 11 is a schematic side view illustrating a configuration of a paint spray tube according to the embodiment of the present invention.
FIG. 12 is a schematic cross-sectional view illustrating a state when the ablator layer according to the embodiment of the present invention is subjected to aerodynamic heating.
FIG. 13 is a cross-sectional view of a shroud portion showing a configuration of the artificial satellite according to the embodiment of the present invention.
FIG. 14 is a cross-sectional view showing a configuration of a solar cell paddle according to the embodiment of the present invention.
FIG. 15 is a schematic diagram for explaining a flow of deployment of the solar cell paddle according to the embodiment of the present invention.
FIG. 16 is a schematic diagram for explaining a flow of deployment of the solar cell paddle according to the embodiment of the present invention.
FIG. 17 is a perspective view showing a state of the artificial satellite according to the embodiment of the present invention during orbit operation.
FIG. 18 is a schematic sectional view showing a configuration of a thrust generating device according to an embodiment of the present invention.
FIG. 19 is a schematic perspective view showing an example of the configuration of the attitude control mechanism in the pitching direction of the artificial satellite according to the embodiment of the present invention.
FIG. 20 is a schematic perspective view showing another example of the configuration of the attitude control mechanism in the pitching direction of the artificial satellite according to the embodiment of the present invention.
[Explanation of symbols]
1 Two-stage rocket
2 Artificial satellite
2a Shroud part
2b Nose part
3 1st stage solid motor
4 2nd stage solid motor
5 Intersection
6 Fairing
6a Stagnation point
7 Low solar absorptivity layer
8 Ablator layer
8a Thermal protection layer
8b Structure layer
8c carbon fiber
9 Thermal insulation layer
10 Metal sheet layer
11-13 Equipment mounting panel
14 Internal equipment
15 Vent
16 Fairing materials
17 Sealing material
17a Masking
18 Female Jig
18a hole
19 Male jig
19a protrusion
20 electromagnet
21 Body
22, 23 Honeycomb structure part
24 Paint supply device
25 Body
26 Paint spray tube
27 props
28 Internal space
29 Masking Department
30 nozzles
32 Masking Department
33 Missing part
34 solar battery paddle
34a hinge part
34b bent part
35,36 solar panel
35a, 36a Cell mounting plate
35b, 36b solar cell
35c radiator
35d Axis
36c heat protection board
37,38 Electric motor
37a, 38a wire
39 Motor controller
40 acceleration sensor
41 Sun sensor
42 Thrust generator
43 cases
44 Piping
45 nozzle
46 heater
47 On-off valve
48a, 48b temperature sensor
49 Pressure sensor
50 phase change substances
51 Sublimation gas
52a, 52b, 53a, 53b Guide tube
52c, 53c wire
52d, 53d weight
54,55 electric motor
56 Earth Sensor
57 Moving weight
58 linear motor
58a Stator
58b runner

Claims (3)

An artificial satellite that travels in orbit with a sparse atmosphere, its outer shell has a streamlined shape that tapers in the direction of travel, and functions as a fairing at launch. . The satellite according to claim 1, wherein the outer shell has a heat-resistant structure against aerodynamic heating. The outer shell has an ablator layer and a low solar absorptivity layer made of a heat-resistant film having a low solar absorptivity that covers the surface of the ablator layer. 3. The artificial satellite according to claim 2, further comprising a ventilation hole through which the pyrolysis gas generated from the ablator layer is passed to be released to the outside.

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