JP3923890B2 - Attitude control device and artificial satellite - Google Patents

Description

[0001]
BACKGROUND OF THE INVENTION
The present invention relates to an attitude control apparatus that performs attitude control of an artificial satellite using gravity inclination and an artificial satellite including the attitude control apparatus.
[0002]
[Prior art]
Artificial satellites such as earth observation satellites and communication satellites need to maintain a certain attitude with respect to the earth. Disturbance torque acts on satellites in orbital operation due to drag from the gas in the upper atmosphere, gravity gradient, interference with geomagnetism, solar radiation pressure, gas leakage, etc. It is necessary to correct the attitude of the artificial satellite disturbed by the disturbance torque.
[0003]
In addition, the artificial satellite includes a solar cell paddle to supply power to the onboard equipment. A solar battery paddle mounted on a conventional artificial satellite has a configuration in which a plurality of solar battery panels are connected in series by hinge coupling, for example, provided on both side surfaces of a hexahedral artificial satellite body. Until the orbit is inserted, it is stored in a bellows-like state on the side of the satellite body. In such a conventional solar cell paddle unfolding structure, the solar cell paddle is biased in the unfolding direction by a spring, and the energizing state is resisted by a holding mechanism such as a hook. It is configured to hold the folded state. When the satellite is put into orbit, the holding mechanism is released and the solar cell panels are deployed so as to be linearly arranged by the biasing force of the spring.
[0004]
In such a conventional solar cell paddle deployment structure, once the deployment of the solar cell paddle is completed, the deployment angle cannot be changed. Therefore, in order to increase the power generation efficiency of the solar cell paddle as much as possible, it is necessary to direct the light receiving surface of the solar cell panel to the sun. At this time, the solar cell paddle cannot be moved alone, and the artificial satellite itself However, it is necessary to perform attitude control so that the light receiving surface faces the sun.
[0005]
For this reason, an attitude control device is mounted on the artificial satellite. Conventional attitude control devices for artificial satellites include devices using thrust generators, devices using magnetic torquers, and devices using gravity gradients.
[0006]
[Problems to be solved by the invention]
However, many of the conventional attitude control devices are housed in the main body of the artificial satellite, and occupy a part of the space for accommodating the artificial satellite, resulting in an increase in size and weight of the artificial satellite. In particular, in an attitude control device using a thrust generator, when the amount of liquid fuel mounted is increased in order to extend the life of the artificial satellite, the artificial satellite becomes larger and heavier.
[0007]
Satellites with a mass of 1000kg or more are classified as large, 500kg to 1000kg are small, 100kg to 500kg are mini, 10kg to 100kg are micro, and 10kg or less are pico. Sometimes. Among such artificial 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 at the target time and target altitude and After being transferred to the track and the fairing being discarded during the transfer, it is put into this track.
[0008]
Once a satellite has started operating in orbit, it can be easily repaired and refurbished like a ground facility even if a failure occurs that makes it impossible to carry out the planned mission. I can't. For this reason, the equipment for artificial satellites is subjected to a thorough test and inspection prior to launch so as to ensure the reliability of the equipment so as not to cause problems as much as possible in orbit.
[0009]
Especially for large and small satellites, the system is complex, the missions are diverse, and the number of devices installed is large. Therefore, many tests and inspections for ensuring reliability are complicated and take a long time. Will be carried out over a long period of time. Also, a lot of tests and inspections are required for the devices and the like used for assembling such artificial satellites.
[0010]
On the other hand, for pico-type and micro-type satellites, the system is relatively simple, many of them are dedicated to a single mission, and the number of equipment is relatively small, so large and small This method has the advantage that the number and period of tests and inspections required are smaller than those of other satellites, and the development period and development cost are reduced.
[0011]
In addition, for example, earth observation satellites and communication satellites are required to shorten their development period as much as possible in order to cope with sudden changes in the global environment and technical obsolescence of equipment mounted on artificial satellites. Therefore, there is an increasing demand for pico-type and micro-type satellites used for these applications.
[0012]
The present invention has been made in view of such circumstances, and it is not necessary to occupy the space for storing the artificial satellite, and the artificial satellite is reduced in size and weight as much as possible. It is an object of the present invention to provide an attitude control device capable of attitude control and an artificial satellite using the attitude control device.
[0013]
[Means for Solving the Problems]
In order to solve the above problems, an attitude control device according to the present invention includes a solar cell paddle pivotally attached to a part of a satellite body, a first actuator attached to the satellite body, and the first actuator. An actuator and a wire connecting the solar cell paddle, and driving the wire by operating the first actuator to change the deployment angle of the solar cell paddle to control the attitude of the artificial satellite by gravity inclination It is characterized by what it should do.
[0014]
The artificial satellite according to the present invention includes an outer shell that functions as a fairing when launched, a solar cell paddle pivotally attached to a part of the outer shell, and a first attached to the inside of the outer shell. An actuator, and a wire connecting the first actuator and the solar cell paddle; the wire is driven by operating the first actuator; and a deployment angle of the solar cell paddle is changed to change a gravitational gradient It is characterized by the fact that the posture control by means of
[0015]
By adopting such a configuration, it is possible to control the attitude of the artificial satellite using the solar cell paddle that has been conventionally used as a power generation system for the artificial satellite, so there is no need to provide a separate attitude control device. Therefore, it is not necessary to divide the space for storing the artificial satellite for the attitude control device.
[0016]
As described above, the artificial satellite is housed in a fairing provided at the head of the rocket in order to protect the artificial satellite from the air resistance received during the launch. In the conventional artificial satellite, an artificial satellite including an outer shell and a fairing are provided separately, and the artificial satellite is housed in the fairing and mounted on the rocket. Therefore, in the artificial satellite according to the present invention as described above, since the outer shell of the artificial satellite functions as a fairing, it is not necessary to provide a fairing separately from the artificial satellite. It is possible to launch using a rocket (for example, a two-stage small rocket). Therefore, in the artificial satellite according to the present invention, the spatial efficiency when mounting the artificial satellite on the rocket used for the launch can be greatly improved as compared with the conventional one, and as a result, the launch vehicle for launching the artificial satellite is obtained. Overall drastic downsizing and cost reduction can be realized.
[0017]
Usually, the mass ratio of the artificial satellite to the whole rocket is about 1%, and the mass ratio of the rocket fuel to the whole rocket is about 90%. Therefore, the miniaturization and weight reduction of the artificial satellite are greatly reduced. And a reduction in the amount of rocket fuel loaded, it is possible to drastically reduce launch costs.
[0018]
Further, when the artificial satellite according to the present invention is configured as a pico-type or micro-type artificial satellite, a launch rocket dedicated to the artificial satellite can be configured. In addition to launching, a pico-type or micro-type satellite can be put into an optimal altitude and orbit.
[0019]
In the above invention, the center of gravity of the artificial satellite is moved in a direction in which the artificial satellite maintains a predetermined posture with respect to the center of the earth by operating the first actuator in accordance with an output signal from the earth sensor. It is desirable to configure as much as possible.
[0020]
Many artificial satellites are provided with an earth sensor, and the position of the artificial satellite is always controlled with respect to the center of the earth by detecting the approximate center position of the earth by the earth sensor. Therefore, with such a configuration, the deployment angle of the solar battery paddle is changed according to the output signal from the earth sensor, and the center of gravity position of the artificial satellite is moved in the direction in which the artificial satellite maintains a predetermined attitude. The attitude control of the artificial satellite can be performed only by moving the solar cell paddle alone, and the energy consumption can be greatly reduced as compared with the prior art.
[0021]
Moreover, in the said invention, two wires pulled out from two places separated in the direction which cross | intersects both the gravitational direction of the said solar cell paddle and the expansion | deployment direction of the said solar cell paddle, and the extraction length of this wire Each of the second actuators, and two weights attached to the respective leading ends of the wires pulled out, and by operating the second actuator, the two lead-out lengths of the wires It is desirable that the position of the artificial satellite be controlled by gravity gradient by changing the magnitude of gravity acting on each of the weights by relative change.
[0022]
Further, in the above invention, a weight capable of moving along a moving direction intersecting both the gravitational direction of the solar cell paddle and the deploying direction of the solar cell paddle, and the weight in the moving direction. A third actuator that moves along the moving direction, and moves the weight along the moving direction to control the attitude of the artificial satellite by gravity gradient. It is desirable.
[0023]
Thereby, it is possible to perform attitude control in a direction that intersects with the deployment direction of the solar cell paddle (for example, the pitching direction when the deployment direction of the solar cell paddle is the rolling direction of the artificial satellite) with a simple configuration. .
[0024]
Moreover, in the said invention, the light-receiving surface of the said solar sensor goes to the sun according to the incident angle to the said light-receiving surface of the sunlight which the solar sensor which has a light-receiving surface parallel to the light-receiving surface of the said solar cell paddle detects Thus, it is desirable that the operation of the first actuator be controlled so as to change the deployment angle of the solar cell paddle.
[0025]
By doing so, the solar cell paddle can be moved independently so that a larger solar radiation intensity is received by the light receiving surface of the solar cell paddle, and the energy consumption is greatly reduced compared to the conventional case. Can do. In addition, since the amount of liquid fuel or the like can be reduced, the satellite itself can be reduced in size and weight.
[0026]
DETAILED DESCRIPTION OF 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 the 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, a first stage solid motor 3, a second stage solid motor 4, and an interstage section 5 according to the present invention. And mainly consists of
[0027]
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. It is comprised from the part 2b (refer FIG. 2). The mass of the artificial satellite 2 is 45 kg, for example, and is classified as a micro artificial satellite. A first stage solid motor 3 is connected below the artificial satellite 2 via an interstage part 5.
[0028]
Further, the interstage part 5 has a cylindrical shape substantially the same diameter as the cylinder of the artificial satellite 2, and a second stage solid motor 4 is supported therein.
[0029]
The first stage solid motor 3 has a movable nozzle, and flight control is performed by changing the injection direction of this nozzle. The second stage solid motor 4 does not have a structure for performing special flight control.
[0030]
The launch process of the artificial satellite 2 is as follows. First, by starting combustion of the first stage solid-state motor 3, the two-stage rocket 1 starts to rise from a range 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 continues to burn for 40 seconds, and at the end of combustion, the flight altitude reaches an altitude of about 45 km. Next, at the end of the combustion of the first stage solid motor 3, that is, after 40 seconds from the start of launch, the first stage solid motor 3 is disconnected as soon as possible, and the second stage solid motor 4 starts to burn. . The second stage solid motor 4 continues to burn for 39 seconds, the flight speed at the end of the combustion reaches a maximum flight speed of about 7.7 km / s, and the flight altitude reaches an altitude of about 240 km. After the combustion of the second stage solid motor 4 is finished, the second stage solid motor 4 and the interstage part 5 are disconnected from the artificial satellite 2 and the artificial satellite 2 is put into the ballistic orbit.
[0031]
Further, another example of the launching process of the artificial satellite 2 is as follows. First, by starting combustion of the first stage solid-state motor 3, the two-stage rocket 1 starts to rise from the launch field 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.
[0032]
About 39 seconds after the start of combustion of the first stage solid motor 3, the two-stage rocket 1 starts a gravity turn eastward at a pitch of about 2 ° / second. The first stage solid motor 3 finishes combustion after 40 seconds from the start of combustion, and thereafter is separated from the second stage rocket 1 as soon as possible.
[0033]
The second-stage solid motor 4 starts burning about 2 seconds after the completion of the combustion of the first-stage solid motor 3. The second stage solid motor 4 finishes combustion after about 39 seconds from the start of combustion. At that 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. . Also, the altitude reached at this time is about 185 km, and the artificial satellite 2 is put into orbit here. If the artificial satellite 2 is further reduced in weight, the reaching altitude can be further increased.
[0034]
Such a two-stage rocket 1 will fly at a significantly higher speed than that of a conventional multi-stage rocket having three or more stages. Therefore, the flight environment such as the heating rate, dynamic pressure, and enthalpy received by the two-stage rocket 1 during flight is much harsher than that received by the multistage rocket having three or more stages. According to the calculation by the inventors of the present application, the maximum aerodynamic heating rate of the two-stage rocket 1 launched from the launching field near the pole and flying ballistically is about 2.6 MW / m. ² The maximum dynamic pressure is about 13 atmospheres, and the maximum acceleration is about 31G. These are 10 times larger than those of large rockets such as H2A rockets developed by the Space Development Agency, and how the flight environment of the two-stage rocket 1 according to this embodiment is You can see if it is harsh.
[0035]
In order to withstand such a severe flight environment, the artificial satellite 2 according to the present embodiment has a configuration as described below. FIG. 2 is a schematic cross-sectional view showing an outline of the 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, and is used as an outer shell of the artificial satellite 2.
[0036]
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 laminated in order from the outside. The low sunlight absorption rate layer 7 uses a white paint having sufficient resistance to aerodynamic heating during the flight of the two-stage rocket 1. Thus, by covering substantially the entire fairing 6 with the low solar absorptivity layer 7 having a low solar absorptance, when the fairing 6 is exposed to sunlight, the solar heat is absorbed. The temperature rise inside the fairing 6 can be suppressed as much as possible.
[0037]
In addition, the white system here is a thing with comparatively high lightness regardless of an achromatic color or a chromatic color. That is, the white system includes, for example, a slightly bluish white color. Moreover, in this Embodiment, although it was set as the structure which makes the low sunlight absorptivity layer 7 a white-type layer, it is not limited to this, 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 solar absorptivity low enough to maintain the internal space of the fairing 6 at a temperature allowed by the equipment provided therein. Needless to say.
[0038]
The ablator layer 8 is made of a material used as an ablator for a rocket such as CFRP. In the present embodiment, the ablator layer 8 is made of CFRP. FIG. 3 is a cross-sectional view showing 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 two layers, a thermal 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-like cloth bodies obtained by cutting carbon fiber woven into a strip shape are embedded in a synthetic resin. It has a structure that can achieve thermal protection performance and keep thermal stress low.
[0039]
On the other hand, the structural layer 8b, which is the inner layer of the ablator layer 8, is composed of CFRP in which a plurality of carbon fibers woven in a continuous cloth shape are laminated, and has high specific strength even if carbonized at high temperatures. In addition, the structure can maintain a high specific rigidity. By adopting such a structure, it is not necessary to provide a metal structure layer on the fairing 6, and it is possible to achieve strength and rigidity comparable to the metal fairing while being lightweight.
[0040]
The heat insulation layer 9 is comprised with the heat insulating material of the inorganic material. As shown in FIG. 2, the ablator layer 8 and the heat insulating layer 9 are configured to be thickest at the stagnation point portion 6 a that is the tip portion of the fairing 6, and become thinner from this point toward the rear. This is because, in the stagnation point portion 6a, the aerodynamic heating at the time of flight is maximized and decreases toward the rear, so that the portion closer to the tip of the fairing 6 is required to have resistance to aerodynamic heating. In addition, the thickness of the ablator layer 8 and the heat insulating layer 9 is reduced toward the rear of the fairing 6, thereby contributing to further weight reduction of the fairing 6.
[0041]
The metal sheet layer 10 is an aluminum sheet.
[0042]
Inside the fairing 6, device mounting panels 11, 12, and 13 are provided. Each of the device mounting panels 11 to 13 includes artificial sensors such as various sensors, an on-board computer, a storage device, and antennas. A satellite internal device 14 is mounted. These internal devices 14 are configured to perform radiant heat exchange with the metal sheet layer 10, whereby the temperature of the internal devices 14 is, for example, an allowable temperature range of a general electronic device from −15 ° C. Properly maintained to be in the range of 45 ° C.
[0043]
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 about 1690 ° C. at the maximum. During the flight of the two-stage rocket 1, the fairing 6 is exposed to such a high temperature, so that pyrolysis gas is generated from the thermal protection layer 8 a, and the temperature increase inside the fairing 6 is caused by the endothermic reaction at this time. In addition to this, the pyrolysis gas covers the surface of the fairing 6 and further heating of the satellite 2 is prevented.
[0044]
Further, when the fairing 6 is subjected to such aerodynamic heating, the thermal protection layer 8a is carbonized and changed to black. Here, when the low solar absorptance layer 7 is not provided, the surface of the fairing 6 changes to black throughout, and the solar satellite 2 absorbs sunlight during orbital operation of the artificial satellite 2, It is conceivable that the internal temperature of the artificial satellite 2 increases abnormally. 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 the low solar absorptivity layer 7 has sufficient resistance to aerodynamic heating as described above. Since the coating material is used, the low solar absorptivity layer 7 does not disappear by aerodynamic heating when the two-stage rocket 1 flies, and the surface of the fairing 6 is maintained white. For this reason, the solar absorptivity of the surface of the artificial satellite 2 can be lowered, and an increase in the internal temperature of the artificial satellite 2 can be suppressed. In addition, it is not necessary to apply a heat resistant paint to a part where the surface temperature is expected to exceed the use upper limit temperature of the heat resistant paint and where the aerodynamic heating is high.
[0045]
FIG. 5 is a partial perspective view of the fairing 6 showing the configuration of the low solar absorptance 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 sunlight absorption rate layer 7 is provided with a large number of air holes 15. That is, the heat protection layer 8a is exposed only on the portion of the vent hole 15 in the surface of the fairing 6. As described above, pyrolysis gas is generated from the thermal protection layer 8a while the fairing 6 is subjected to aerodynamic heating. As shown in FIG. 6, this pyrolysis gas passes through the vent hole 15 and is released to the outside of the artificial satellite 2. On the other hand, when the vent hole 15 is not provided in the low solar absorptivity layer 7, there is no escape route for the pyrolysis gas generated from the thermal protection layer 8a due to aerodynamic heating, and the low solar absorptivity Cracks or delamination will occur in the layer 7. Therefore, by forming the air holes 15 in the low solar absorptivity layer 7, the occurrence of cracks or peeling of the low solar absorptivity layer 7 can be suppressed.
[0046]
That is, by making the fairing 6 as described above, the heat generated by aerodynamic heating during the flight of the two-stage rocket 1 is prevented from entering the space of the artificial satellite 2 and the orbit operation of the artificial satellite 2 is performed. It is possible to release the heat generated from the internal device 14 to the outside of the aircraft at the time, and to keep the temperature of the internal device 14 at the time of the flight of the two-stage rocket 1 and the orbit operation of the artificial satellite 2. It is possible.
[0047]
In the present embodiment, the diameter of the vent hole 15 is 1 mm, and the distance between two adjacent vent holes 15 is 25 mm. In addition, the ventilation hole 15 makes the exposed area of the thermal protection layer 8a as small as possible, for example, the diameter of the ventilation hole 15 is 3 mm, and the interval between two adjacent ventilation holes 15 is 5 mm. It goes without saying that other dimensions may be adopted as long as the desired performance of releasing the pyrolysis gas to the outside is exhibited.
[0048]
Next, an example of a method for manufacturing the fairing 6 according to the embodiment of the present invention will be described. FIG. 7 is a schematic diagram for explaining an example of a pre-process for painting the fairing 6, wherein (a) shows the first stage, (b) shows the second stage, and (c) shows the third stage. (D) shows the fourth stage. First, a fairing material 16 that 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. And the coating material (for example, white type | system | group) which has a low sunlight absorption factor is apply | coated to the surface of the fairing material 16 as follows.
[0049]
As shown in FIG. 7 (a), in the first step of the pre-painting process of 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 applied. It is put on a female jig 18 which is manufactured according to the outer shape of the fairing 6 in advance. In the female jig 18, a large number of holes 18a are provided in advance in accordance with the positions of the air holes 15 provided in the low sunlight absorption layer 7 of the fairing 6.
[0050]
Next, in the second stage, as shown in FIG. 7B, the female jig 18 is placed on the fairing material 16.
[0051]
In the third stage, as shown in FIG. 7 (c), a male jig 19 manufactured in advance according to the outer shape of the fairing 6 is put on the sealing material 17. A large number of protrusions 19 a are provided on the inner surface of the male jig 19 in advance in accordance with the positions of the air holes 15 provided in the low sunlight absorption rate layer 7 of the fairing 6. Then, alignment is performed so that the protrusion 19 a is inserted into the hole 18 a, and the male jig 19 is pressed against the fairing material 16. At this time, the portion of the sealing material 17 that overlaps the protrusion 19a passes through the hole 18a and comes into contact with the surface of the fairing material 16, and the sealing material 17 is adhered only at a position corresponding to the vent hole 15 of the fairing material 16. It is done.
[0052]
In the fourth stage, as shown in FIG. 7 (d), 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 large number of sticking portions of the fairing material 16 of the sealing material 17 remain as masking (masking portions) 17a on the surface of the fairing material 16, and this portion of the ablator layer 8 is hidden. Thus, the pre-process for painting the fairing 6 is completed.
[0053]
A heat resistant paint having a low solar absorptance 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.
[0054]
Next, another example of the method for manufacturing the fairing 6 according to the embodiment of the present invention will be described. FIG. 8 is a schematic diagram for explaining another example of a pre-process for painting the fairing 6, wherein (a) shows the first stage and (b) shows the second stage.
[0055]
As shown in FIG. 8 (a), in the first stage of the pre-painting process of the fairing 6 in this example, the electromagnet 20 that has been manufactured according to the inner shape of the fairing 6 is placed inside the fairing material 16. Deploy.
[0056]
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 has a main body portion 21 having a substantially similar shape obtained by reducing the fairing material 16 and having a plurality of coils (not shown) provided therein, and an outer side and an inner side of the main body portion 21, respectively. The honeycomb structure portions 22 and 23 are formed. 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. The outer surface of this will be in close contact over substantially the whole.
[0057]
The honeycomb structure portions 22 and 23 are made of metal, and have a structure in which a plurality of hexagonal holes penetrates in the thickness direction, and a cocoon is also formed in a honeycomb shape. The honeycomb structure portions 22 and 23 can make the electromagnet 20 lightweight and highly rigid.
[0058]
Further, the coil provided in the main body portion 21 has the thickness direction of the main body portion 21 as a central axis, and by applying a direct current to the coil, for example, as shown in FIG. It is possible to generate a magnetic field with N as the N pole and the inside as the S pole.
[0059]
In the present embodiment, the electromagnet 20 is arranged inside the fairing material 16, but the present invention is not limited to this, and a permanent magnet may be arranged instead of the electromagnet 20. Moreover, although the electromagnet 20 is configured to have the honeycomb structure portions 22 and 23, the present invention is not limited to this. The electromagnet 20 is made of a metal or metal and a synthetic resin having substantially the same outer shape as the inner shape of the fairing material 16. It is good also as a structure which accommodated the some coil in the housing made from a composite material.
[0060]
Next, in the second stage of this example, as shown in FIG. 8 (b), a coating material supply device 24 whose inside is manufactured in advance so as to roughly match the outer shape of the fairing 6 is put on the outside of the fairing material 16.
[0061]
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 coating material supply device 24 is formed of a flexible material and has a bag-like body portion 25 larger than the fairing material 16 and a plurality of coating materials protruding from a plurality of locations on the inner surface of the body portion 25. It is mainly composed of an injection tube 26 and a support column 27 protruding from a plurality of locations different from the protruding location of the paint spray tube 26 on the inner surface of the main body portion 25.
[0062]
The main body portion 25 has an internal space 28, and a heat-resistant paint having a low sunlight absorption rate supplied from the outside to the internal space 28 can be passed therethrough. The main body portion 25 has flexibility as described above, and is formed in a shape that substantially follows the outer shape of the fairing material 16. Therefore, when the fairing material 16 is covered with the coating material supply device 24, the main body portion 25 is deformed according to the outer shape of the fairing material 16, and the main body portion 25 is covered over substantially the entire fairing material 16. Is possible.
[0063]
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. In addition, 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 blocked by the masking portion 29, the heat resistant paint is not ejected from the other end.
[0064]
Further, a support column 27 made of a high magnetic permeability material is projected from the inner surface of the main body portion 25. One end of the column 27 is fixed to the inner surface of the main body portion 25, and a masking portion 32 is attached to the other end. The masking portion 32 is also a permanent magnet. For example, when the electromagnet 20 generates a magnetic field with the N pole on the outside of the main body portion 21 and the S pole on the inner side, the masking portions 29 and 32 have the N pole on the close side of the main body portion 25 and the S side on the remote side. It is attached to the paint spray tube 26 and the column 27 so as to generate a magnetic field as a pole.
[0065]
When covering the fairing material 16 with the coating material supply device 24 in the second stage, the masking portions 29 and 32 are placed on the outer surface of the fairing material 16 where the vent 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 this magnetic field. Thus, the pre-process for painting the fairing 6 is completed.
[0066]
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 sprayed from the nozzle 30. The nozzle 30 is provided so as to incline so that the heat-resistant paint is sprayed toward the surface of the fairing material 16, and thereby the portions on which 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 excluding.
[0067]
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. Thereby, the fairing 6 is completed.
[0068]
As described above, by constructing the artificial satellite 2 in which the fairing 6 is integrated as an outer shell, a fairing is prepared separately from the artificial satellite like the conventional artificial satellite, and the artificial satellite is combined with the fairing. Eliminates the need for mounting on a rocket. An artificial satellite requires an outer shell structure, and in order to ensure sufficient strength and rigidity of the entire artificial satellite, a strength member or the like is required in the case of a conventional artificial satellite. In the artificial satellite 2 according to the present embodiment, the fairing 6 also serves as an outer shell structure, and the fairing 6 has sufficient strength and rigidity as described above, so that it is necessary to provide a strength member or the like separately. Absent.
[0069]
Further, as a result of the trial calculation by the inventors of the present application, the fairing 6 according to the embodiment of the present invention mainly composed of CFRP having a specific gravity smaller than that of a metal is a conventional artificial satellite equipped with the same internal device. Compared to the metal fairing used when launching the rocket with a three-stage rocket, it is understood that both have the same mass when configured to ensure the same strength and rigidity. It was. That is, the artificial satellite 2 according to the present invention requires the conventional artificial satellite when the internal devices mounted on the artificial satellite are the same as the total mass of the conventional artificial satellite and the metal fairing. It is possible to reduce the weight by the mass of the outer shell structure and the strength member. For example, when the mass of a 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. Therefore, the artificial satellite 2 according to the present embodiment equipped with the same internal device is used. Then, the weight can be reduced by 15 kg from the total mass of the conventional satellite and fairing.
[0070]
By reducing the weight of such an artificial satellite (and fairing), the rocket itself that transports it to outer space can be made smaller. In addition, the structure can be greatly simplified by constructing the two-stage rocket 1 instead of the conventional three-stage or more multi-stage rocket as in the present embodiment, so that the development cost and the development period can be greatly reduced. Etc. can be greatly reduced.
[0071]
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 as the rocket configured to separate and release the fairing. Using a rocket, it is possible to launch an artificial satellite 2 having a mass about 15 kg larger than that of the artificial satellite. Paradoxically speaking, in a satellite integrated with a fairing of the same mass as that of a satellite having a structure for separating and releasing a conventional fairing, it can be launched with a significantly smaller rocket.
[0072]
Further, as described above, the two-stage rocket 1 flies at a significantly higher speed than a conventional multi-stage rocket having three or more stages. For example, when a gravity 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 becomes 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 the second-stage rocket 1 (artificial satellite 2) at the time of the release when the artificial satellite 2 is released 10 seconds later. ) Is about 260 km.
[0073]
Thus, the horizontal movement distance from the start of launch of the artificial satellite 2 to the time of release is about 260 km, and this distance is extremely shorter than the horizontal move distance from the start of launch of the conventional artificial satellite to the time of release. . Therefore, when launching an artificial satellite with a conventional three-stage rocket, for example, after launching from a launch site in Japan, it was sometimes impossible to observe the launch from the launch of the satellite to the release by an observation station 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 Tanegashima Space Center launch site, the flight status of the two-stage rocket 1 until the release of the artificial satellite 2 is shown. And Ogasawara Pursuit Station, and no foreign facilities are required for observation. As described above, since the observation from the launch of the two-stage rocket 1 to the launch of the orbit of the artificial satellite 2 can be performed with only the facilities in Japan, the launch work of the rocket is made extremely simpler than before. be able to.
[0074]
Further, 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 time of release is observed from the launch of the two-stage rocket 1 to the launch of the artificial satellite 2 by one observation station. It is also a distance that can be done. Therefore, if the observation from the two-stage rocket 1 to the orbit insertion of the artificial satellite 2 is consistently observed at one observation station, the launch operation of the rocket can be further simplified and the cost can be further increased. In addition, it can be expected to reduce labor.
[0075]
As described above, in the two-stage rocket 1 according to the second embodiment, the ablator layer 8 of the fairing 6 is carbonized by aerodynamic heating. At this time, carbon fibers remain in a state where the carbon fibers overlap each other in the vicinity of the surface of the ablator layer 8. FIG. 12 is a schematic cross-sectional view for explaining a state when aerodynamic heating is applied to the ablator layer 8 according to the embodiment of the present invention. The thermal protection layer 8 a of the ablator layer 8 is carbonized at the initial stage of launching the two-stage rocket 1. As shown in FIG. 12, the carbon fibers 8c that have been carbonized and remain in the thermal protection layer 8a overlap in a state where the longitudinal direction of the fibers is directed in various directions. The binding force of each carbon fiber 8c is relatively weak, and when the force is applied from the outside, the carbon fiber 8c moves slightly so as to be shifted from each other.
[0076]
Here, as described above, the two-stage rocket 1 according to the present embodiment receives an extremely high dynamic pressure as compared with the conventional one at the time of flight. Therefore, 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.
[0077]
However, when the carbon fibers 8b (and the carbon fibers remaining in this layer when the structural layer 8c is carbonized) move minutely to each other three-dimensionally, the ablator layer 8 functions as a seismic isolation structure. The vibration applied to the surface of 6 is prevented from being transmitted to the inside as it is. Thereby, it becomes difficult for external vibration to be transmitted to the internal device 14, and occurrence of a failure of the internal device 14 can be suppressed.
[0078]
Further, when the artificial satellite 2 is put into orbit, the second stage solid-state motor 4 and the interstage portion 5 are separated from the artificial satellite 2, but at this time, an impact is applied to the artificial satellite 2. This impact is also mitigated inside the artificial satellite 2 by the above-described seismic isolation function of the fairing 6, and the occurrence of a failure of the internal device 14 is suppressed.
[0079]
Next, the configuration of the artificial satellite 2 will be described in more detail. FIG. 13 is a cross-sectional view of shroud portion 2a showing the configuration of artificial satellite 2 according to the embodiment of the present invention. As shown in FIG. 13, in the portion from the rear end of the shroud portion 2a of the artificial satellite 2 to the front of a predetermined length, a part of the fairing 6 in the circumferential direction is missing, so that the missing portion 33 is blocked. An unfolded solar cell paddle 34 is provided.
[0080]
The solar cell paddle 34 has a configuration in which two plate-like solar cell 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 attached to one end of the missing portion 33. Thereby, the solar cell panel 35 can be rotated outward from the state where the missing portion 33 is closed, with the pivot portion as a center.
[0081]
The other end of the solar cell panel 35 is hinged to one end of the solar cell panel 36 by a hinge portion 34a. Before the two-stage rocket 1 is launched, the solar cell panels 35 and 36 are accommodated so as to close the missing portion 33 in a superposed state.
[0082]
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 side) solar cell panel 35 has a structure in which solar cells 35b are attached to one surface of a cell attachment plate 35a. Further, a radiator 35c is provided so as to close the missing portion 33, and the radiator 35c functions as a partition wall between the artificial satellite 2 and outer space. The radiator 35c discharges the heat generated by the internal device 14 to the outside. The optical characteristics and area of the radiator 35 c are set by the amount of heat generated by the internal device 14. Further, when the solar cell paddle 34 is stored, the other surface of the cell mounting plate 35a faces the outer surface of the radiator 35c, and the solar cell 35b is disposed on the outer surface.
[0083]
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. If exhaust heat is unnecessary and is small enough, a partition made of a composite material such as CFRP may be provided instead of the radiator 35c.
[0084]
The other (front end side) solar cell panel 36 has a cell mounting plate 36a. When the solar cell paddle 34 is stored, the solar cell 36b is opposed to the solar cell 35b. It is attached to one side of 36a. That is, the solar battery cell 36b is attached to the inner surface of the cell attachment plate 36a. Further, a heat protection plate 36c composed of an ablator layer, a heat insulating layer, and the like is attached to the other surface (surface outside the machine) of the cell attachment plate 36a. As a result, it is possible to exert a thermal protection function against aerodynamic heating received when the two-stage rocket 1 flies, and to prevent the temperature of the internal device 14 from being excessively increased. Further, when the solar battery paddle 34 is housed, the thermal protection plate 36c is positioned at the outermost part, so that the solar battery cells 35b and 36b are protected from the aerodynamic heating received when the two-stage rocket 1 flies.
[0085]
As shown in FIG. 13, two electric motors 37 and 38 are provided inside the artificial satellite 2. One end of a wire 37 a is connected to the output shaft of one electric motor 37, and the other end of the wire 37 a is connected to a pivot 35 d that pivotally attaches 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. The winding direction of the wire 37a around the pivot 35d on the other end is such that when the wire 37a is drawn from the pivot 35d, the solar cell panel 35 is rotated outward. The direction of movement is the direction in which torque is generated (that is, the counterclockwise direction in FIG. 13). As a result, 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 shaft 35d, and is rotated in the direction rotating to the solar cell panel 35. Torque will act.
[0086]
On the other hand, the one end side of the wire 38a is similarly wound around the output shaft of the electric motor 38. The other end of the wire 38a is connected to the hinge portion 34a. When the solar cell paddle 34 is housed in the fairing 6, the wire 38a is fixed by the output shaft of the electric motor 38 so that a predetermined tension acts. Alternatively, the solar battery paddle 34 may be locked to the fairing 6 so as not to expand by the lock mechanism, and the lock mechanism may be released when the artificial satellite 2 is put into orbit.
[0087]
Further, inside the artificial satellite 2, a motor controller 39 for controlling the operations of the electric motors 37 and 38 and an acceleration sensor 40 connected to the motor controller 39 are provided.
[0088]
In addition, a spring (not shown) is provided between the solar cell panels 35 and 36, and the solar cell panels 35 and 36 are urged by the springs in a direction in which they are developed. Further, a stopper (not shown) is provided between the solar cell panels 35 and 36, and when the solar cell paddle 34 is stored in the fairing 6, the solar cell panel 35, 36 is maintained in an overlapped state against the biasing force of the spring.
[0089]
Next, a deployment operation of the solar cell paddle 34 will be described. FIG.15 and FIG.16 is a schematic diagram for demonstrating the flow of expansion | deployment of the solar cell paddle 34 which concerns on embodiment of this 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 part 5 are separated 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 electrically operates 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 about the pivot 35d toward the outside of the apparatus. Further, when the solar cell paddle 34 is rotated, the missing portion 33 is closed only by the radiator 35c.
[0090]
As shown in FIG. 15, the motor controller 39 drives the electric motors 37 and 38 until the solar cell paddle 34 rotates just 90 °. When the solar cell paddle 34 is rotated 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. Will rotate in a direction away from the solar cell panel 35 (see FIG. 16). When the solar cell panel 36 further rotates due to the inertial force and the solar cell panel 36 rotates 180 ° and the solar cell panels 35 and 36 are linearly aligned, the solar cell panels 35 and 36 are rotated. A stopper (not shown) provided in a proximity portion of the solar cell panel is engaged with the solar cell panel 35, thereby maintaining the state in which the solar cell panels 35 and 36 are linearly arranged. Thereby, the expansion | deployment operation | work of the solar cell paddle 34 is completed.
[0091]
As described above, since the electric motor 37.38 and the solar cell paddle 34 are connected by the light wires 37a and 38a and the electric motor 37 and 38 are driven, the solar cell paddle 34 is 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 cell paddle 34 using a mechanism such as the above.
[0092]
FIG. 17 is a perspective view showing a state during the orbit operation of the 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 battery cell 36 b of the solar battery paddle 34. The sun sensor 41 outputs an electrical signal corresponding to the solar radiation intensity, and the output destination is the motor controller 39. The configuration of the solar sensor 41 will be described in more detail. A so-called coarse sensor (CSS) is used for the solar sensor 41. This sensor makes sunlight incident through a slit, and the incident angle of the sunlight. It is the structure which measures only the component of an orthogonal direction to this slit, and is a well-known thing. The solar sensor 41 is arranged so that the light receiving surface of the solar cell paddle 34 and the light receiving surface thereof are parallel to each other. The motor controller 39 drives the electric motors 37 and 38 so that the light receiving surface of the solar sensor 41 faces the sun, and thereby the solar cell paddle 34 so that the light receiving surface of the solar cell paddle 34 faces the sun. The unfolding angle 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, there are few cases where the light receiving surface is completely opposed. The deployment angle of the solar cell paddle 34 is controlled so as to face each other. The solar cells 35b and 36b receive solar radiation and the earth albedo (earth reflection component of solar radiation) to generate power, and the electric power obtained by the solar cells 35b and 36b is supplied to the internal device 14. It has become.
[0093]
When the mass of the artificial satellite 2 is 45 kg, the required power can be estimated to be about 20 W from the statistical result on the relationship between the past 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-side orbit, when the day-side orbit time and the night-side orbit time are the same, the required power generation amount is about 40 W, which is twice the required power. . There are two types of solar cells that have abundant use results in space: an expensive but highly efficient gallium arsenide type and a silicon type that is inexpensive but has relatively low power generation efficiency. If silicon type solar cells are used as the solar cells 35b, 36b, the light receiving surface area required for 80 W power generation, which is a value obtained by multiplying the required power generation amount by 2 as a safety factor, is about 0.6 m. ² It becomes. As a result of the inventors' calculation, the area of the solar battery paddle 34 of the artificial satellite 2 having a mass of 45 kg is 0.9 m. ² Therefore, even if silicon type solar cells are used, it is possible to obtain a necessary amount of electric power.
[0094]
When the artificial satellite 2 is put into orbit, the solar cell paddle 34 is deployed and, if necessary, an S-band antenna or a gravitational tilt beam accommodated in the artificial satellite 2 is deployed (not shown) to artificially The orbit insertion of the satellite 2 is completed.
[0095]
As shown in FIG. 17, the artificial satellite 2 travels toward the tip of the nose portion 2b. However, a slight amount of air exists in the space around the artificial satellite 2, and the artificial satellite 2 has an air resistance. Will receive. When the satellite body has a hexahedron or cylindrical shape like a conventional satellite, the influence of such air resistance is large, and the satellite decelerates, and the balance between centrifugal force and gravity is It 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 as the artificial satellite body has a streamline shape, it is less susceptible to air resistance than the conventional artificial satellite. The deceleration due to is suppressed as much as possible.
[0096]
However, the artificial satellite 2 according to the present invention continues to receive some air resistance, decelerates, and the altitude drop needs to be raised by the thrust generated by the thrust generator. Conventional large-sized, small-sized, or mini-type artificial satellites use a thrust generating device configured to burn liquid fuel or to inject liquid nitrogen or the like. When the target operational life of the artificial satellite is set to 1 year or longer, a large amount of fuel or injection material is required to maintain the altitude of the artificial satellite. In addition, such a thrust generator has a complicated structure, and the size and mass of the device itself are large. Therefore, a large-sized, small-sized, or mini-type artificial satellite has a relatively large equipment mounting volume, and it is relatively easy to mount such a conventional thrust generator, but the artificial satellite 2 according to the present embodiment is similar to the artificial satellite 2 according to the present embodiment. A micro-type or pico-type satellite has a small equipment mounting volume, and it is not easy to mount these thrust generators on a micro-type or pico-type satellite.
[0097]
Therefore, the artificial satellite according to the present embodiment is equipped with a thrust generator as described below. FIG. 18 is a schematic cross-sectional view showing the configuration of the thrust generator 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. Case 43 has a rectangular parallelepiped box shape, and phase change material 50 is accommodated therein. As the phase change material 50, a material having a melting point as low as about 300K such as eicosane is used.
[0098]
One surface of the case 43 is opened, and a pipe 44 is extended from the opening. The end of the pipe 44 is connected to the apex portion of the nozzle 45 that is expanded in a substantially conical shape. In addition, an opening / closing valve 47 is provided in the middle of the pipe 44, and when the opening / closing valve 47 is closed, a portion from the connection position of the case 43 and the pipe 44 to the case 43 to the opening / closing valve 47. Constitutes a closed space, and when the on-off valve 47 is opened, the space in the case 43 communicates with the outside through the pipe 44 and the nozzle 45.
[0099]
Further, a heater 46 is provided so as to cover substantially the entire outer peripheral surface of the case 43, the pipe 44 and the nozzle 45. The heater 46 is of a type that converts electric power such as a heating wire into heat, and the case 43, the pipe 44, and the nozzle 45 are heated by receiving the supply of electric power from the outside.
[0100]
Further, 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 the middle of the nozzle 45 and the on-off valve 47 of the pipe 44.
[0101]
Next, the operation of the thrust generator 42 according to the present embodiment will be described. If it is determined by the on-board computer (not shown) built in the artificial satellite 2 that thrust is required for the artificial satellite 2, electric power is supplied to the heater 46 from the outside, and the case 43, the piping 44, and The nozzle 45 is heated. The temperature of the internal space of case 43 rises, sublimation gas 51 of phase change material 50 is generated (increased), and the internal pressure of 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 the power supply 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.
[0102]
After the sublimation gas 51 is injected, 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 brought to a predetermined low temperature side allowable temperature or less due to adiabatic expansion of the gas. If necessary, these are heated by a heater 46 so as not to be rapidly cooled. Further, after the sublimation gas 51 is injected, it is necessary for the on-board computer to close the on-off valve 47 and monitor the output value of the temperature sensor 48a to control the temperature of the phase change material 50 within an appropriate range. For example, the case 43 is heated by the heater 46.
[0103]
Next, the result of trial calculation for the necessary amount of the phase change material 50 will be described. In the following description, the cross-sectional area viewed from the traveling direction is 0.7 m. ² The result of the trial calculation when the altitude of the artificial satellite 2 is set to an altitude of 350 km where the air resistance is relatively large is shown. According to the calculation of the present inventor, when the mass of the artificial satellite is 45 kg, the deceleration per day can be canceled by the ejection of 15 g of the sublimation gas 51. In this case, when the operational life of the artificial satellite 2 is 100 days, the required mass of the phase change material 50 is 1.5 kg. For example, when eicosane is used for the phase change material 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.
[0104]
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 several tens of kg at most, this The purpose used to maintain the altitude and speed of the artificial satellite 2 can sufficiently achieve its purpose.
[0105]
Next, a configuration related to attitude control during orbit 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. Square tubular guide tubes 52a and 52b are provided along the front end edges of the solar cell panels 35 and 36 at the front end of the solar cell panels 35 and 36 (downstream side end in the traveling direction of the satellite 2), respectively. Similarly, guide tubes 53a and 53b are respectively provided along the rear end edges of the solar cell panels 35 and 36 at the rear end of the solar cell panels 35 and 36 (upstream side end in the traveling direction of the artificial satellite 2). . As shown in FIG. 19, the guide tubes 52a and 52b are arranged so as to be linearly arranged when the solar cell paddle 34 is deployed. Similarly, the guide tubes 53a and 53b are also solar cells. The paddles 34 are arranged linearly when the paddles 34 are unfolded. Wires 52c are inserted into the linearly arranged guide tubes 52a and 52b, and the inner end of the wire 52c is attached to the output shaft of the electric motor 54 provided in the spacecraft 2. Yes. A weight 52d is attached to the machine end of the wire 52c.
[0106]
Similarly, a wire 53c is inserted into the linearly arranged guide tubes 53a and 53b, and the inner end of the wire 53c is connected to the output shaft of the electric motor 55 provided in the spacecraft 2. It is attached. Further, a weight 53d is attached to the outer end of the wire 53c.
[0107]
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 controlled by the motor controller. As shown in FIG. 17, an earth sensor 56 is provided so as to protrude from the artificial satellite 2 toward the earth side, and the motor controller is connected to the earth sensor 56.
[0108]
The earth sensor 56 outputs an electrical signal corresponding to each of the roll angle and the pitch angle of the artificial satellite 2 with respect to the earth. Of these, the motor controller receives an output signal representing the pitch angle, and the motor controller The operation of the electric motors 54 and 55 is controlled in order to adjust the drawing length of the wires 52c and 53c. More specifically, for example, proportional control is performed so that the pitch angle of the artificial 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 side. By controlling the operation of the electric motors 54 and 55, 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 more gravity from the earth than the weight 53d. As a result, the center of gravity of the artificial satellite 2 moves forward, the leading end of the artificial satellite 2 approaches the earth, and the trailing end is The attitude control of the artificial satellite 2 in the pitching direction is performed so as to be away 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 tip side of the artificial satellite is separated from the earth, Further, the attitude control of the artificial satellite 2 in the pitching direction is performed so that the rear end side approaches the earth.
[0109]
And it is comprised so that operation control of the electric motors 54 and 55 may be performed so that the attitude | position of the pitching direction of the artificial satellite 2 may be maintained so that the earth sensor 56 may always face the center of the earth.
[0110]
The guide tubes 52a, 52b, 53a, 53b can also be used as strength members of the solar cell paddle 34.
[0111]
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 this example, as shown in FIG. 20, a moving weight 57 is provided on the outer edge of the solar cell paddle 34 on the outside of the artificial satellite 2, that is, the opposite edge of the hinge portion 34a of the solar cell panel 36. Yes. The moving weight 57 is attached to the traveling element 58b of a linear motor 58 having a stator 58a and a traveling element 58b provided over the entire length of the edge of the solar cell panel 36, and is attached to the stator 58a. Along with this, it is possible to move integrally with the traveling element 58b.
[0112]
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 related to the pitch angle of the artificial satellite 2 from the earth sensor 56. ing. Also here, the operation of the linear motor 58 is controlled by a known control method as described above. When the moving weight 57 moves in the front-rear direction of the artificial satellite 2, 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.
[0113]
In addition, it is good also as a structure which provides the movement weight which can move to the front-back direction inside the fairing 6. FIG.
[0114]
Further, the attitude control of the artificial satellite 2 in the rolling direction is performed by adjusting the deployment angle of the solar cell 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 representing the roll angle of the artificial satellite 2 from the earth sensor 56, and the motor controller 39 sets the detected roll angle of the artificial satellite 2 to a predetermined target roll angle (this In the case of the embodiment, the rotation direction and the rotation speed of the electric motors 37 and 38 using a known control method such as proportional control, PI control, PID control, or fuzzy control so as to approach the roll angle of 0 °). To control. At this time, the solar battery paddle 34 is tilted to the opposite side of the direction in which the artificial satellite 2 is inclined with respect to the center of the earth, thereby moving the lateral center of gravity position of the artificial satellite 2, resulting in a roll angle of 0 °. The attitude control in the rolling direction of the artificial satellite 2 is performed by tilting the artificial satellite 2 in a direction approaching to.
[0115]
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 cell paddle 34 is adjusted so that the light receiving surface of the solar cell paddle 34 faces the sun. Be controlled. Thereby, the solar cell paddle 34 can be operated independently and the light receiving surface thereof can be directed to the sun, and it is necessary to control the attitude of the artificial satellite itself in order to direct the light receiving surface of the solar cell paddle to the sun as in the prior art. There is no.
[0116]
【The invention's effect】
In the case of the attitude control device according to the present invention and an artificial satellite using the same, the attitude control of the artificial satellite can be performed using a solar cell paddle that has been conventionally used as a power generation system for an artificial satellite. There is no need to provide a separate attitude control device, and therefore there is almost no need to divide the satellite storage space for the attitude control device.
[0117]
Further, since the outer shell of the artificial satellite functions 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 the optimum size for the artificial satellite. Therefore, in the artificial satellite according to the present invention, the spatial efficiency when mounting the artificial satellite on the rocket used for the launch can be greatly improved as compared with the conventional one, and as a result, the launch vehicle for launching the artificial satellite is obtained. Overall drastic downsizing and cost reduction can be realized.
[0118]
In addition, the deployment angle of the solar cell paddle is changed according to the output signal from the earth sensor, and the center of gravity position of the artificial satellite is moved in the direction in which the artificial satellite maintains a predetermined attitude. Therefore, the solar cell paddle is moved alone. The present invention has excellent effects such as the attitude control of the artificial satellite can be performed by itself and the energy consumption can be greatly reduced as compared with the conventional one.
[Brief description of 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.
FIG. 2 is a schematic cross-sectional view showing an outline of the configuration of the artificial satellite according to the embodiment of the present invention.
FIG. 3 is a cross-sectional view showing a configuration of an ablator layer according to an embodiment of the present invention.
FIG. 4 is a graph showing changes in fairing surface temperature, metal sheet layer temperature, and internal device temperature during flight of a two-stage rocket according to an embodiment of the present invention.
FIG. 5 is a partial perspective view of a fairing showing a configuration of a low solar absorptance layer according to an embodiment of the present invention.
FIG. 6 is a schematic partial cross-sectional view of a fairing according to an embodiment of the present invention.
FIGS. 7A and 7B are schematic diagrams for explaining an example of a pre-process for painting a fairing according to an embodiment of the present invention, in which FIG. 7A shows the first stage, FIG. 7B shows the second stage, 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 for explaining another example of a pre-process for painting a fairing according to an embodiment of the present invention, in which FIG. 8A shows the first stage and FIG. 8B shows the second stage. It is a schematic diagram shown respectively.
FIG. 9 is a partial cross-sectional view showing a configuration of an electromagnet according to an embodiment of the present invention.
FIG. 10 is a partial perspective view showing the configuration of the paint supply apparatus according to the embodiment of the present invention.
FIG. 11 is a schematic side view showing the configuration of the 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 an embodiment of the present invention is subjected to aerodynamic heating.
FIG. 13 is a cross-sectional view of the shroud portion showing the 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 an embodiment of the present invention.
FIG. 15 is a schematic diagram for explaining a development flow of the solar cell paddle according to the embodiment of the present invention.
FIG. 16 is a schematic diagram for explaining a development flow of the solar cell paddle according to the embodiment of the present invention.
FIG. 17 is a perspective view showing a state during the orbit operation of the artificial satellite according to the embodiment of the present invention.
FIG. 18 is a schematic cross-sectional view showing a configuration of a thrust generator 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 First stage solid motor
4 Second stage solid motor
5 Steps
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 Panels with equipment
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 part
22, 23 Honeycomb structure part
24 Paint supply device
25 Body part
26 Paint spray tube
27 prop
28 interior space
29 Masking part
30 nozzles
32 Masking part
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 cells
35c radiator
35d Axis
36c thermal barrier
37,38 Electric motor
37a, 38a wire
39 Motor controller
40 Acceleration sensor
41 Solar sensor
42 Thrust generator
43 cases
44 Piping
45 nozzles
46 Heater
47 On-off valve
48a, 48b Temperature sensor
49 Pressure sensor
50 Phase change material
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 (8)

A solar battery paddle pivotally attached to a part of the artificial satellite body and rotated around an axis along the advancing direction of the artificial satellite to change a deployment angle;
A first actuator attached to the satellite body;
A wire connecting the first actuator and the solar cell paddle;
And the weight that can be moved in the traveling direction of the satellite along the outboard edge of the artificial satellite of the solar arrays,
And a third actuator for moving the weight to the traveling direction,
The wire is driven by operating said first actuator, said varying the deployment angle of the solar arrays, also the weight is moved to the traveling direction by operating the third actuator An attitude control device characterized in that the attitude control of the artificial satellite is performed by gravity inclination. The attitude control in the rolling direction of the artificial satellite is performed by changing the deployment angle of the solar cell paddle, and the attitude control in the pitching direction of the artificial satellite is performed by moving the weight in the traveling direction. The attitude control device according to claim 1. By operating the first actuator according to an output signal from an earth sensor mounted on the artificial satellite, the position of the center of gravity of the artificial satellite in a direction in which the artificial satellite maintains a predetermined attitude with respect to the center of the earth attitude control device according to claim 1 or 2, characterized in that are no in order to move. In accordance with the incident angle of sunlight to the light receiving surface detected by a solar sensor having a light receiving surface parallel to the light receiving surface of the solar cell paddle, the solar cell paddle is arranged so that the light receiving surface of the solar sensor faces the sun. to vary the deployment angle, the posture control device according to any one of claims 1 to 3, characterized in that it is configured to control the operation of the first actuator. An outer shell that functions as a fairing during launch,
Is pivotally mounted to a portion of the outer shell, the solar arrays to change the deployment angle by being rotated about an axis along the direction of travel of artificial satellites,
A first actuator mounted inside the outer shell;
A wire connecting the first actuator and the solar cell paddle;
A weight capable of moving in the traveling direction of the artificial satellite along the outer edge of the artificial satellite of the solar cell paddle;
And a third actuator for moving the weight to the traveling direction,
The wire is driven by operating said first actuator, said varying the deployment angle of the solar arrays, also the weight is moved to the traveling direction by operating the third actuator An artificial satellite characterized by attitude control by gravity tilt. The attitude control in the rolling direction of the artificial satellite is performed by changing the deployment angle of the solar cell paddle, and the attitude control in the pitching direction of the artificial satellite is performed by moving the weight in the traveling direction. The artificial satellite according to claim 5. By operating the first actuator in accordance with an output signal from an earth sensor attached to the outer shell, the center of gravity position of the artificial satellite is set in a direction in which the outer shell maintains a predetermined attitude with respect to the center of the earth. The artificial satellite according to claim 5 or 6 , wherein the artificial satellite is intended to move. In accordance with the incident angle of sunlight to the light receiving surface detected by a solar sensor having a light receiving surface parallel to the light receiving surface of the solar cell paddle, the solar cell paddle is arranged so that the light receiving surface of the solar sensor faces the sun. The artificial satellite according to any one of claims 5 to 7 , wherein an operation of the first actuator is controlled so as to change a deployment angle.

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