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Advanced spacecraft design for asteroid sample return mission
Pergamon
Acta Astronautica Vol. 35. Suppl., pp. 231-238, 1995 Elsevier Science Ltd. Printed in Great Britain
0094·5765(94)00188·X
ADVANCED SPACECRAFT DESIGN FOR ASTEROID SAMPLE RETURN MISSION Ichiro Nakatani, Masashi Hashimoto, Hirobumi Saito, Institute of Space and Astronautical Science Masayuki Kamimura and Masaki Adachi Space Systems Development Division, NEC Corporation
ABSTRACT A new generation of M-series launch vehicle, M- V, is currently being developed at Institute of Space and Astronautical Science and the first flight is scheduled for 1996. We will carry out several planetary missions using M- V which can launch a spacecraft of a few hundred kilograms into a trans-planet orbit. Lunar penetorator mission "LUNAR-A" and Mars orbiter mission "PLANET- B" have been approved and the launch is scheduled for 1997 and 1998, respectively. Our follow-on planetary missions will be more advanced and complex because of the increasingly demanding scientific requirements. One of them is the asteroid sample return mission. In order to realize advanced planetary missions, it is absolutely necessary to establish highly advanced and light-weight spacecraft design under the constraint of launch capability. In this paper, we will describe our advanced spacecraft design concept for the asteroid sample return mission with M- V vehicle as an example.
INTRODUCTION The Institute of Space and Astronautical Science (lSAS) is currently developing a new type of launch vehicle, "M- V', the first flight of which is scheduled for 1996. The new vehicle M- V will be able to send payload of 200 - 550kg into transplanetary orbit, depending on the target planet and launching year. Let us classify our planetary missions into six generations, taking into account the degree of mission complexity, as follows: Generation- I :Flyby Mission such as our SAKIGAKE and SUISEI (I) for Halley's Comet in 1990. Generation - II : Orbiter Mission such as our PLANET- B (2) for Mars in 1998. Generation-III: Penetrator or Entry Probe Mission such as our LUNAR-A (3) for lunar penetrator in 1997, and our proposed Venus Balloon Mission. Generation- IV: Lander Mission such as our proposed Mars Rover Mission. Generation - V : Sample Return Mission such as the Asteroid "Nereus" Mission that this paper deals with. Generation- VI: Manned Planetary Mission. Figure 1 shows the dry spacecraft weight vs. mission cost excluding the launch cost for planetary missions of ISAS, as well as NASA as a reference. Note that ISAS has been heading for more complex missions from Generation-I to II and III, under the constraints of spacecraft weight as well as cost. Now we are proposing the asteroid sample return mission belonging to more complex Generation- V. Then it is necessary for ISAS to reduce the spacecraft weight and 231
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Low-cost planetary missions
increase its performance. This point is the main subject of this paper. It is very different from the situation where the spacecraft weight and mission cost of the conventional NASA planetary missions have increased significantly as mission generation evolved from I to Il.Ill and IV ( See figure 1). It, however, resulted in reducing the mission frequency in NASA planetary missions. At present, a new trend of medium size through microsize spacecraft is emerging in the planetary missions >( VB: of the United States. For example, Clementine mission in S; 1994 and NEAR mission, one of the Discovery missions, in NASNUSA GLL 1000 VGR • 1997 utilize light- weight onboard instruments including ~ PPO M:' • •MGN miniature optical sensors, thrusters and high performance ~ • M73 : M71 ·MO computers, which were developed directly or influenced by DISCOVERY o. -G..... t!•• h DoD's intercepter programs. 100 ;67-·IU-~::rol;•• IJ: . J .. .~. ¢ Orbiter Meanwhile, apanese space commumties do not A O+-GeDeratioD 111: have such a strong spring board as the military technolo0 ISAS/JAPAN Pe•• 'rol.r,Prebe M .i•• o X-GeaentioD IV: gies. Therefore ISAS independently comes up with our La.d.r o own strategy for the weight reduction as well as high 10 100 1000 performance of spacecraft. For this purpose, we have SPACECRAFT DRY MASS (ke) initiated the "STRAIGHT" project (STudy on the Reduction Figurel Mission Cost and Spacecraft of Advanced Instrument weiGHT) to develop the advanced, Dry Mass for Various Planetary light-weight spacecraft hardware. The basic concept is Planetary Mission Generations described in the section below. Some of the advanced spacecraft technologies are described in "ADVANCED TECHNOLOGIES" section. In "ASTEROID SAMPLE RETURN MISSION" section, the asteroid "Nereus" sample return is referred to as a model mission for our next generation mission to apply our advanced spacecraft technologies. The final section is a conclusion.
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BASIC CONCEPT ON WEIGHT REDUCTION OF ADVANCED SPACECRAFT When spacecraft sizing is discussed, it is essential to describe spacecraft functions in terms of information, energy and mechanics. Information and energy are collected into and re-emitted from spacecraft through its surface (area-dependent function). Handling of the information and the energy can be intensively performed inside the spacecraft. Mechanical subsystems such as structure subsystem and propulsion subsystem are highly dependent on the spacecraft weight. These are conceptually illustrated in figures 2 and 3.
,t
.......
e
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Baerl)' Hudlia,
§
Req.......t
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Id'o~.tie.
HaDdUa,
Structure ',
_.
··.u
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Figure 2 Conceptual Diagram of Spacecraft Function based on Information, Energy, Mechanics, and Area-dependent Function
Figure 3 Conceptual Flow Chart of Spacecraft Weight Reduction
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(a) Information Handling : Information handling is included in many subsystems such as power / communication / command and data handling / attitude and orbit control/thermal control. Information handling would be the most integration-oriented process in spacecraft. The recent development in microelectronics provides light-weight electronics onboard instruments with low power consumption such as high density mass memories (memory density O.l-lGbitslkg) and high speed processors (several 10MIPS with several kg and watts). These figures of merit would increase even more in the near future by orders of magnitude. The objective of the missions in the generations I to IV is information aquisition with appropriate information processing that follows. In addition, planetary missions require autonomous operations in space. Reduction of the weight as well as power consumption of the equipment for those operations will result in the reduction of the weights of the power, and mechanical subsystems. Therefore, information handling system would be the most powerful driver for advanced, light- weight spacecraft. (b) Energy Handling: Power subsystem (except for solar paddles), and thermal control subsystem (except for radiator panels) are included in this item. Electrical power drives all electrical circuits. It also converts to thermal energy in heaters and to kinetic energy in electrical propulsion or momentum wheels. Energy, however, would not be so integration-oriented as information. For example, the stored energy density would improve by a factor (not by orders of magnitude) as the secondary batteries evolve from conventional Ni- Cd (""30Whlkg) to Ni- MH (""50Whlkg), or more advanced Li-ion type (""2OOWhlkg). In parallel with the technology advancement in the increase of the energy (or energy flow) density, power consumption as well as heater power has to decrease as much as possible. (c) Receiving / Transmitting of Information and Energy (Area-Dependent Function) : Receiving / transmitting of information and energy depend on the surface area. It means that a spacecraft has to expose a certain surface area to outside space, even though the surface area could be decreased by technological improvements. In communications, a certain area of antenna aperture is required to satisfy the communication link to the earth station. The area can be decreased by increasing RF power and data compression rate, using sophisticated coding techniques etc. Communications with higher frequency such as Ka band or optical light would be very effective to decrease the area. Optical, radiowave, and particle sensors to collect information on the space environment, planets or fixed stars require a certain surface area in order to keep their sensitivity or their resolution. The area can be decreased by the improvement of their detector's sensitivity or arraying technique. Total electric power to be generated at a solar paddle is proportional to its area. At 1A.D. distance from the sun, electric power of 140-200 W/rn2 can be generated with realistic conversion efficiency. This electric power drives all the electronic circuits and then converts to thermal energy, which have to be radiated through radiator panels. A typical value of heat radiation density is 350 W/m2 for T=300' K, e =0.76 . Reduction of total power consumption is very effective to reduce both area of the solar paddle and the radiator panel. (d) Mechanics (Mass-dependent Function) : Structure, propulsion and attitude control actuators are highly dependent on the total spacecraft mass. Structure subsystem, in most cases, has its weight - percentage of 15%- 25% in the total spacecraft weight. Also, in planetary missions, the dry mass of propulsion subsystem and the propellant are considerably large. Size of the propulsion system and amount of the propellant to be used are approximately proportional to the spacecraft mass. On the other hand, the structure would be subject to the "Elastic Scaling" (4) , in which the dimension normal to the main load (diameter) scales in proportion to the 3/2 power of the length against buckling and bending for the acceleration at launch. It means that half-size spacecraft needs only sixteen times less massive structure weight compared with the original.
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The basic concept for spacecraft weight reduction is conceptually described in figure 3. Power requirement determines the weight of energy handling subsystem approximately by a rate of lO-20Wlkg. The antennas jand sensors with a certain area of aperture are required by communications and observation performance. Also there is some weight for information handling. The weight of mechanical subsystems (structure / propulsion) is approximately in proportion to the subtotal weight of these three functions. The weight and power in each function could be reduced by the advance of technologies. These effects propagate sequentially to down stream. Let us take an example of our generation-II spacecraft such as PLANET-B. In the generation - II spacecraft, the weight percentage of each function is approximately as follows: Information handling-10%, energy handling-10%, sensor and antenna-5%, structure and propulsion-75%. As described before, the information handling would be most integrationoriented. Assume that we successfully reduce the weight of the information handling down to 1/3 in the use of advanced technologies of micro-electronics and micro-packaging. This result in reduction of the total spacecraft weight down to 73% Furthermore, it seems possible to reduce the power rewuirement, antenna and sensor size by a factor of 20%, and to improve a figure of weight-power ratio of the energy handling by a factor of 20%. These additional improvements result in reduction of the total spacecraft weight down to 55%. Although the situation of spacecraft system in reality is more complex, the above-mentioned concept seems to be qualitatively reasonable. Based upon this basic concept, details of the advanced technologies to be applied is described in the next section.
ADVANCED TECHNOLOGIES We have initiated the "STRAIGHT' project to develop the advanced, light-weight spacecraft hardware, as described in the introduction. Based on the activities of STRAIGHT, our advanced design of the future planetary spacecraft is described as follows. (a) Information handling : In the past, the functions of spacecraft bus system have been allocated to each subsystem (ex. power / communications / command and data handling / attitude and orbit control/thermal control) mainly for their electrical functions. Each subsystem has performed collection and analysis of necessary information and control on its components, respectively. We are planning to integrate the functions of information collection, analysis and component control in a space bus synthetic management system called as "Bus Management Unit". This concept is very effective to enhance the information handling capabilities including autonomous operations as well as to reduce the weight with micro packagingtechnologies. The system block diagram of the Bus Management Unit (BMU) is shown in figure 4. In the BMU, there are three "managers", one memory bank for their common use and network system for their linkage. All onboard judgement will be performed by these managers, and telemetry from / commands to components will be basically transferred through external data bus. Data to be stored in data recorder or to be transmitted to the ground will flow along the dedicated line, and "data handling manager" will control the data flow with "telemetry data interface". "Autonomous manager" will break the requirements from ground down to several operation units and will estimate their availabilities based on time and power resources. The operation unit sequence will be controlled by "sequence controller" in the autonomous manager and the controller will send requirement to other two managers. For the mission having no requirement of autonomous operations, the function of autonomous manager will be reduced just to "resource monitoring and control". The data handling manager and "attitude / orbit control manager" will break the requirements from the autonomous manager down to several commands for each component, and control each component through the external data bus. Each manager will collect information from components to write them and sequence status in the common memory bank. Information transmission between three managers will be performed through the data bank. If a manager become out of use, other two managers will replace it with the information in this bank.
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BMU structural configuration is shown in figure 5. Its structure consists of several stacked palettes. One palette consists of one or two print wiring board assemblies and structural frames with internal connectors.. Each frame will be made by FRP ,magnesium or aluminum. One palette is connected to neiboring palettes with some elastomeric internal connectors in its frame. This connectorless instrumentation method will be applied basically to all the components to reduce the component size. Weight reduction trend related to BMU functions is shown in figure 6. For MUSES-B, which will be launched in 1996, BMU functions are dispersing to 9 units (Battery Charge Control Unit, Power Control Unit, Command Signal Selector, CoMmand Decoder, Data Handling Unit, House Keeping unit, Attitude Orbit Control Electronics, Attitude Orbit COntrol Processor and Heater Control Electronics) and these total weight is about 40kg. We are developing the components for PLANET - B in 1998. The weight for BMU functions in PLANET- B spacecraft has already reduced to 18kg. Our target 9kg will be realized as a flight model at the biginning of 21st century. (b) Energy Handling: High frequency switching regulator, NiMH battery cell (stored enegy density 50Whlkg) and high efficiency Si solar cell (efficiency 18%) will be applied to above - mentioned PLANET - B. Furthermore, we are investigating the development of piezoelectric type transformer, Li-ion battery (stored energy density""'200Wh/kg) for the planetary spacecraft of the next generation. (c) Receiving / Transmitting of Information and Energy (Area-Dependent Function) : We are planning to develop the small, light and high efficiency power amplifier and transponder for downlink modulated by PSK method. These improvements would lead to a decrease in the parabolic antenna diameter. Also, light weight parabolic antenna with meshed FRP is under development. In the other planetary missions like the asteroid sample return, the solar intensity gets low. Hence, it becomes essencial to develop a light-weight, high efficiency solar paddle. This includes such technologies as the flexible CulnSe2 solar cell, the light-weight rigid paddle, and the inflatable solar paddle. As a thermal transmission device, micro heat pipe is under development. (d) Mechanics (Mass-Dependent Function) : In PLANET- B, composite material will be fully utilized. Especially, the high pressure tank of He pressurant will be made from carbon fiber reinforced plastics overwrapped on a thin Ti liner. The weigh of the composite tank is about half of a conventional Ti tank. We will keep exploiting advanced technologies of composite materials in the next generations. ASTEROID SAMPLE RETURN MISSION MISSION SCENARIO Let us descibe the asteroid nereus sample return mission as a model mission for the next generation planetary missions. The spacecraft with 365kg wet mass will be launched into the transfer orbit for the Nereus in januarY,2002 by the 5-staged M- V vehicles. Instead of swing-by technique or chemical propulsion, ion thruster will be used to accelerate continuously the spacecraft to the Nereus with the low thrust level of 7.5mN. The specific impulse and electrical consumption power are 3,120sec and 250W, respectively. The spacecraft will approach the Nereus by controlling the orbit and attitude using the reaction control system and optical images of the onboard camera as well as the laser-radar (LIDAR). At the altitude of 10 meters from the Nereus surface, the spacecraft will drop the visual target to the surface. Then the relative position and attitude between the spacecraft and the visual target will be determined for the final approach to the
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surface. These processes should be performed based on the highly autonomous onboard functions. Staying at 1 meter altitude, the spacecraft will take asteroid samples of 1 kg to the sample container. A solid motor will be,used to depart for the earth in November, 2003. The spacecraft will return to the earth in January, 2006. The small reenter capsule containing the sample will be released from the spacecraft to reentry into the earth atomosphere. The capsule will be decelerated by the parachute and then recovered on the ground.
Figure 4 BMU System Block Diagram ..,- - - - - - - -, - - _. ... -.-' 0"'11 SU'PLY MODULI
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Figure 5 BMU Structural Configuration
Figure 6 Weight Reduction related to BMU Functions
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APPLICATION OF STRAIGHT PROJECT We performed the case study of this asteroid sample return mission based on the STRAIGHT project. This mission belongs to the generation- V planetary mission and requires highly autonomous operations due to the large communication delay and the low communication data rate . Under the severe constraint of spacecraft weight, the information handling should be performed effectively, in the full use of information on spacecraft itself and the asteroid collected by various sensors. BMU with only 9kg weight plays a role in all the information handling in the spacecraft. The power requirement is relatively large (350W at 2 A.U.) because the electrical propulsion is used in the mission. It is very effective to increase the performance of the solar paddles. The light- weight rigid solar paddle with 40W/kg, which is higher by a factor of than conventional one, will be developed to the mission. Also, the refined NiMH battery with 50Whlkg will be used The advancement of communications subsystem include light-weight CFRP parabolic high gain antenna, the light-weight S- band transponder, and the X-band power amplifier with 30% conversion efficiency. Weight and power budget for the asteroid sample return mission is shown in table 1 on the basis of above-mentioned technologies. The reasonable system margin (64.9kg,21.6%) is shown on the table 1. All these advanced technologies make the mission feasible with enough margin of resources. CONCLUSIONS ISAS will perform several planetary missions with our M- V vehicle which is currently being developed. These planetary missions evolve from relatively simple flyby missions to more sophisticated missions such as the asteroid sample return, which this paper deals with. In order to perform these advanced planetary missions under severe constraint of spacecraft weight, the advanced technologies should be exploited to develop light- weight spacecraft with high performance. We have initiated the "STRAIGHT" project for this purpose. The spacecraft function would be devided into four functions; information handling, energy handling, area -dependent function (aperture antenna or sensor), mass-dependent mechanical function. Then we proposed the conceptual flow chart of spacecraft weight reduction. Based on the basic concept, we discussed our present strategy for spacecraft weight reduction as follows. First priority is to reduce weight of the information handling subsystem down to 1/3 or less. In parallel, the other subsystems (power requirement, energy handling, antenna/sensor, and structure/propulsion) do their best to reduce the weight or to improve the performance by a factor of 20% or more. These advance of spacecraft technologies would lead to reduction of the spacecraft weight in half.We proposed the Bus Management Unit as an integrated information handling instrument. Several key technologies in the other three functions are described in the paper. We performed the case study to apply these advanced technologies to the asteroid sample return mission. It is confirmed that these advanced technologies make the asteroid sample return mission feasible even under severe constraint of spacecraft weight. REFERENCE (1) (2) (3) (4)
K.Hirao and T.Itoh : "Project Overview and Highlights of Suisei and Sakigake", Advances in Space Research, Vo1.5, No.12, 1985, pp.55-64 . I.Nakatani et al. : "Planet - B: Japanese Mars Mission", IAF-93-Q.3.400, IAF, Graz, Austria (1993). H.Mizutani et al. : "Lunar Interior Exploration by Lunar Penetrator Mission", IAF 90-039, IAF, Dresden , GDR (1990). K.Aaron: "Elastic Scaling of Small Structures", 7th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA (1993).
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Table 1 Weight and Power Budget Table Subsystem/Component
Weight (kll:)
Power Solar Cell Panel Baterry Shunt Dissipator Communicat ion High Gain Antenna 2-axis Gimbal System S-band Transponder S-band Diplexer S-band Switch X-band Transmitter X-band Power Amp. Low Gain Antenna Data Handling Bus Management Unit Data Recorder Launch Ope ration Attittxle/Orbit Control Moment mil Wheel 2-dim. Sun Sensor Opt ical Nav. Cam era 2-axis Gimbal Syst em Laser Altimeter Star Tracker Rate Gyro Assy. Accelerometer Propulsion Reaction Control System Ret urn Tri p Motor Electric Propultion Stntcture Integration Mechanical Thermal Wireharness Balance Weight Mission Payload Sampler Target Plate Re- entry Capsule Scientific Instrument Ory We4!ht Total Fuel ( N2 H4IXe) System Margin Total Weil!'ht Heater Total Power Generation Power @ 2AU • EP
Electric P ropulsion
40 .0
sr0.0
Power (W) Ep· off on 0.0
35.0 2.2 2.8 15.7
190/0Si 40W/kg NiMH 50WWkg 65 .4
20 .4
q, 1.5m
7.2 4.0 2 0.6 0.1 1.2 0.2 0.4 10.0
15.0 22.8
Tx-Power 5W / Eff.=30%
6.0 0.0
11.0 16.6
Tx-Power 5W / Eff.=30Oh
27.5 1.7
132Mbits
29 .2 27 .5 1.3
0.0 16.7
13.7 76 .9
250.0 19.0 45 .0 25.0
24.6 30.3
22.0 2Nms x 4 3.5 6.0 2 unit (Wide!Te le) 10.0 17.0 6.0 10.0 2 unit 2.4 4 unit
7.2 3.5 0.0 0.0 0.0 6.0 0.0 0.0
8.0 1.5 1.0 1.5 2.0 6.0 0.7 0.4 89.0
Meshed FRP
0.0 14.4
28.8 9.0 1.0
4.0 2 1.1
Remarks
70 .0 0.0
70.0
250.0
0.0
IN x 12Thrulters 382m/s 4.9mN
0.0
0.0 11.5 8.8 7.0 3.0
27 .0
0.0 5.0 1.0 15.0 6.0
TBD TBO TBO
0.0 0.0 0.0 0.0
TBO
259.7 50.0 55.3
N2H4 17Om/slXe 1600m1s
365.0 20 .0 335 .9
20 .0 268.4
350.0
1.4kW @ IAU
Acta Astronautica Vol. 35. Suppl., pp. 231-238, 1995 Elsevier Science Ltd. Printed in Great Britain
0094·5765(94)00188·X
ADVANCED SPACECRAFT DESIGN FOR ASTEROID SAMPLE RETURN MISSION Ichiro Nakatani, Masashi Hashimoto, Hirobumi Saito, Institute of Space and Astronautical Science Masayuki Kamimura and Masaki Adachi Space Systems Development Division, NEC Corporation
ABSTRACT A new generation of M-series launch vehicle, M- V, is currently being developed at Institute of Space and Astronautical Science and the first flight is scheduled for 1996. We will carry out several planetary missions using M- V which can launch a spacecraft of a few hundred kilograms into a trans-planet orbit. Lunar penetorator mission "LUNAR-A" and Mars orbiter mission "PLANET- B" have been approved and the launch is scheduled for 1997 and 1998, respectively. Our follow-on planetary missions will be more advanced and complex because of the increasingly demanding scientific requirements. One of them is the asteroid sample return mission. In order to realize advanced planetary missions, it is absolutely necessary to establish highly advanced and light-weight spacecraft design under the constraint of launch capability. In this paper, we will describe our advanced spacecraft design concept for the asteroid sample return mission with M- V vehicle as an example.
INTRODUCTION The Institute of Space and Astronautical Science (lSAS) is currently developing a new type of launch vehicle, "M- V', the first flight of which is scheduled for 1996. The new vehicle M- V will be able to send payload of 200 - 550kg into transplanetary orbit, depending on the target planet and launching year. Let us classify our planetary missions into six generations, taking into account the degree of mission complexity, as follows: Generation- I :Flyby Mission such as our SAKIGAKE and SUISEI (I) for Halley's Comet in 1990. Generation - II : Orbiter Mission such as our PLANET- B (2) for Mars in 1998. Generation-III: Penetrator or Entry Probe Mission such as our LUNAR-A (3) for lunar penetrator in 1997, and our proposed Venus Balloon Mission. Generation- IV: Lander Mission such as our proposed Mars Rover Mission. Generation - V : Sample Return Mission such as the Asteroid "Nereus" Mission that this paper deals with. Generation- VI: Manned Planetary Mission. Figure 1 shows the dry spacecraft weight vs. mission cost excluding the launch cost for planetary missions of ISAS, as well as NASA as a reference. Note that ISAS has been heading for more complex missions from Generation-I to II and III, under the constraints of spacecraft weight as well as cost. Now we are proposing the asteroid sample return mission belonging to more complex Generation- V. Then it is necessary for ISAS to reduce the spacecraft weight and 231
232
Low-cost planetary missions
increase its performance. This point is the main subject of this paper. It is very different from the situation where the spacecraft weight and mission cost of the conventional NASA planetary missions have increased significantly as mission generation evolved from I to Il.Ill and IV ( See figure 1). It, however, resulted in reducing the mission frequency in NASA planetary missions. At present, a new trend of medium size through microsize spacecraft is emerging in the planetary missions >( VB: of the United States. For example, Clementine mission in S; 1994 and NEAR mission, one of the Discovery missions, in NASNUSA GLL 1000 VGR • 1997 utilize light- weight onboard instruments including ~ PPO M:' • •MGN miniature optical sensors, thrusters and high performance ~ • M73 : M71 ·MO computers, which were developed directly or influenced by DISCOVERY o. -G..... t!•• h DoD's intercepter programs. 100 ;67-·IU-~::rol;•• IJ: . J .. .~. ¢ Orbiter Meanwhile, apanese space commumties do not A O+-GeDeratioD 111: have such a strong spring board as the military technolo0 ISAS/JAPAN Pe•• 'rol.r,Prebe M .i•• o X-GeaentioD IV: gies. Therefore ISAS independently comes up with our La.d.r o own strategy for the weight reduction as well as high 10 100 1000 performance of spacecraft. For this purpose, we have SPACECRAFT DRY MASS (ke) initiated the "STRAIGHT" project (STudy on the Reduction Figurel Mission Cost and Spacecraft of Advanced Instrument weiGHT) to develop the advanced, Dry Mass for Various Planetary light-weight spacecraft hardware. The basic concept is Planetary Mission Generations described in the section below. Some of the advanced spacecraft technologies are described in "ADVANCED TECHNOLOGIES" section. In "ASTEROID SAMPLE RETURN MISSION" section, the asteroid "Nereus" sample return is referred to as a model mission for our next generation mission to apply our advanced spacecraft technologies. The final section is a conclusion.
e
g
-1- J
BASIC CONCEPT ON WEIGHT REDUCTION OF ADVANCED SPACECRAFT When spacecraft sizing is discussed, it is essential to describe spacecraft functions in terms of information, energy and mechanics. Information and energy are collected into and re-emitted from spacecraft through its surface (area-dependent function). Handling of the information and the energy can be intensively performed inside the spacecraft. Mechanical subsystems such as structure subsystem and propulsion subsystem are highly dependent on the spacecraft weight. These are conceptually illustrated in figures 2 and 3.
,t
.......
e
1'ben.al 8.fO
uor/Allta
Baerl)' Hudlia,
§
Req.......t
Prop"I'~D
Id'o~.tie.
HaDdUa,
Structure ',
_.
··.u
·
Figure 2 Conceptual Diagram of Spacecraft Function based on Information, Energy, Mechanics, and Area-dependent Function
Figure 3 Conceptual Flow Chart of Spacecraft Weight Reduction
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(a) Information Handling : Information handling is included in many subsystems such as power / communication / command and data handling / attitude and orbit control/thermal control. Information handling would be the most integration-oriented process in spacecraft. The recent development in microelectronics provides light-weight electronics onboard instruments with low power consumption such as high density mass memories (memory density O.l-lGbitslkg) and high speed processors (several 10MIPS with several kg and watts). These figures of merit would increase even more in the near future by orders of magnitude. The objective of the missions in the generations I to IV is information aquisition with appropriate information processing that follows. In addition, planetary missions require autonomous operations in space. Reduction of the weight as well as power consumption of the equipment for those operations will result in the reduction of the weights of the power, and mechanical subsystems. Therefore, information handling system would be the most powerful driver for advanced, light- weight spacecraft. (b) Energy Handling: Power subsystem (except for solar paddles), and thermal control subsystem (except for radiator panels) are included in this item. Electrical power drives all electrical circuits. It also converts to thermal energy in heaters and to kinetic energy in electrical propulsion or momentum wheels. Energy, however, would not be so integration-oriented as information. For example, the stored energy density would improve by a factor (not by orders of magnitude) as the secondary batteries evolve from conventional Ni- Cd (""30Whlkg) to Ni- MH (""50Whlkg), or more advanced Li-ion type (""2OOWhlkg). In parallel with the technology advancement in the increase of the energy (or energy flow) density, power consumption as well as heater power has to decrease as much as possible. (c) Receiving / Transmitting of Information and Energy (Area-Dependent Function) : Receiving / transmitting of information and energy depend on the surface area. It means that a spacecraft has to expose a certain surface area to outside space, even though the surface area could be decreased by technological improvements. In communications, a certain area of antenna aperture is required to satisfy the communication link to the earth station. The area can be decreased by increasing RF power and data compression rate, using sophisticated coding techniques etc. Communications with higher frequency such as Ka band or optical light would be very effective to decrease the area. Optical, radiowave, and particle sensors to collect information on the space environment, planets or fixed stars require a certain surface area in order to keep their sensitivity or their resolution. The area can be decreased by the improvement of their detector's sensitivity or arraying technique. Total electric power to be generated at a solar paddle is proportional to its area. At 1A.D. distance from the sun, electric power of 140-200 W/rn2 can be generated with realistic conversion efficiency. This electric power drives all the electronic circuits and then converts to thermal energy, which have to be radiated through radiator panels. A typical value of heat radiation density is 350 W/m2 for T=300' K, e =0.76 . Reduction of total power consumption is very effective to reduce both area of the solar paddle and the radiator panel. (d) Mechanics (Mass-dependent Function) : Structure, propulsion and attitude control actuators are highly dependent on the total spacecraft mass. Structure subsystem, in most cases, has its weight - percentage of 15%- 25% in the total spacecraft weight. Also, in planetary missions, the dry mass of propulsion subsystem and the propellant are considerably large. Size of the propulsion system and amount of the propellant to be used are approximately proportional to the spacecraft mass. On the other hand, the structure would be subject to the "Elastic Scaling" (4) , in which the dimension normal to the main load (diameter) scales in proportion to the 3/2 power of the length against buckling and bending for the acceleration at launch. It means that half-size spacecraft needs only sixteen times less massive structure weight compared with the original.
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Low-cost planetary missions
The basic concept for spacecraft weight reduction is conceptually described in figure 3. Power requirement determines the weight of energy handling subsystem approximately by a rate of lO-20Wlkg. The antennas jand sensors with a certain area of aperture are required by communications and observation performance. Also there is some weight for information handling. The weight of mechanical subsystems (structure / propulsion) is approximately in proportion to the subtotal weight of these three functions. The weight and power in each function could be reduced by the advance of technologies. These effects propagate sequentially to down stream. Let us take an example of our generation-II spacecraft such as PLANET-B. In the generation - II spacecraft, the weight percentage of each function is approximately as follows: Information handling-10%, energy handling-10%, sensor and antenna-5%, structure and propulsion-75%. As described before, the information handling would be most integrationoriented. Assume that we successfully reduce the weight of the information handling down to 1/3 in the use of advanced technologies of micro-electronics and micro-packaging. This result in reduction of the total spacecraft weight down to 73% Furthermore, it seems possible to reduce the power rewuirement, antenna and sensor size by a factor of 20%, and to improve a figure of weight-power ratio of the energy handling by a factor of 20%. These additional improvements result in reduction of the total spacecraft weight down to 55%. Although the situation of spacecraft system in reality is more complex, the above-mentioned concept seems to be qualitatively reasonable. Based upon this basic concept, details of the advanced technologies to be applied is described in the next section.
ADVANCED TECHNOLOGIES We have initiated the "STRAIGHT' project to develop the advanced, light-weight spacecraft hardware, as described in the introduction. Based on the activities of STRAIGHT, our advanced design of the future planetary spacecraft is described as follows. (a) Information handling : In the past, the functions of spacecraft bus system have been allocated to each subsystem (ex. power / communications / command and data handling / attitude and orbit control/thermal control) mainly for their electrical functions. Each subsystem has performed collection and analysis of necessary information and control on its components, respectively. We are planning to integrate the functions of information collection, analysis and component control in a space bus synthetic management system called as "Bus Management Unit". This concept is very effective to enhance the information handling capabilities including autonomous operations as well as to reduce the weight with micro packagingtechnologies. The system block diagram of the Bus Management Unit (BMU) is shown in figure 4. In the BMU, there are three "managers", one memory bank for their common use and network system for their linkage. All onboard judgement will be performed by these managers, and telemetry from / commands to components will be basically transferred through external data bus. Data to be stored in data recorder or to be transmitted to the ground will flow along the dedicated line, and "data handling manager" will control the data flow with "telemetry data interface". "Autonomous manager" will break the requirements from ground down to several operation units and will estimate their availabilities based on time and power resources. The operation unit sequence will be controlled by "sequence controller" in the autonomous manager and the controller will send requirement to other two managers. For the mission having no requirement of autonomous operations, the function of autonomous manager will be reduced just to "resource monitoring and control". The data handling manager and "attitude / orbit control manager" will break the requirements from the autonomous manager down to several commands for each component, and control each component through the external data bus. Each manager will collect information from components to write them and sequence status in the common memory bank. Information transmission between three managers will be performed through the data bank. If a manager become out of use, other two managers will replace it with the information in this bank.
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BMU structural configuration is shown in figure 5. Its structure consists of several stacked palettes. One palette consists of one or two print wiring board assemblies and structural frames with internal connectors.. Each frame will be made by FRP ,magnesium or aluminum. One palette is connected to neiboring palettes with some elastomeric internal connectors in its frame. This connectorless instrumentation method will be applied basically to all the components to reduce the component size. Weight reduction trend related to BMU functions is shown in figure 6. For MUSES-B, which will be launched in 1996, BMU functions are dispersing to 9 units (Battery Charge Control Unit, Power Control Unit, Command Signal Selector, CoMmand Decoder, Data Handling Unit, House Keeping unit, Attitude Orbit Control Electronics, Attitude Orbit COntrol Processor and Heater Control Electronics) and these total weight is about 40kg. We are developing the components for PLANET - B in 1998. The weight for BMU functions in PLANET- B spacecraft has already reduced to 18kg. Our target 9kg will be realized as a flight model at the biginning of 21st century. (b) Energy Handling: High frequency switching regulator, NiMH battery cell (stored enegy density 50Whlkg) and high efficiency Si solar cell (efficiency 18%) will be applied to above - mentioned PLANET - B. Furthermore, we are investigating the development of piezoelectric type transformer, Li-ion battery (stored energy density""'200Wh/kg) for the planetary spacecraft of the next generation. (c) Receiving / Transmitting of Information and Energy (Area-Dependent Function) : We are planning to develop the small, light and high efficiency power amplifier and transponder for downlink modulated by PSK method. These improvements would lead to a decrease in the parabolic antenna diameter. Also, light weight parabolic antenna with meshed FRP is under development. In the other planetary missions like the asteroid sample return, the solar intensity gets low. Hence, it becomes essencial to develop a light-weight, high efficiency solar paddle. This includes such technologies as the flexible CulnSe2 solar cell, the light-weight rigid paddle, and the inflatable solar paddle. As a thermal transmission device, micro heat pipe is under development. (d) Mechanics (Mass-Dependent Function) : In PLANET- B, composite material will be fully utilized. Especially, the high pressure tank of He pressurant will be made from carbon fiber reinforced plastics overwrapped on a thin Ti liner. The weigh of the composite tank is about half of a conventional Ti tank. We will keep exploiting advanced technologies of composite materials in the next generations. ASTEROID SAMPLE RETURN MISSION MISSION SCENARIO Let us descibe the asteroid nereus sample return mission as a model mission for the next generation planetary missions. The spacecraft with 365kg wet mass will be launched into the transfer orbit for the Nereus in januarY,2002 by the 5-staged M- V vehicles. Instead of swing-by technique or chemical propulsion, ion thruster will be used to accelerate continuously the spacecraft to the Nereus with the low thrust level of 7.5mN. The specific impulse and electrical consumption power are 3,120sec and 250W, respectively. The spacecraft will approach the Nereus by controlling the orbit and attitude using the reaction control system and optical images of the onboard camera as well as the laser-radar (LIDAR). At the altitude of 10 meters from the Nereus surface, the spacecraft will drop the visual target to the surface. Then the relative position and attitude between the spacecraft and the visual target will be determined for the final approach to the
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surface. These processes should be performed based on the highly autonomous onboard functions. Staying at 1 meter altitude, the spacecraft will take asteroid samples of 1 kg to the sample container. A solid motor will be,used to depart for the earth in November, 2003. The spacecraft will return to the earth in January, 2006. The small reenter capsule containing the sample will be released from the spacecraft to reentry into the earth atomosphere. The capsule will be decelerated by the parachute and then recovered on the ground.
Figure 4 BMU System Block Diagram ..,- - - - - - - -, - - _. ... -.-' 0"'11 SU'PLY MODULI
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Figure 5 BMU Structural Configuration
Figure 6 Weight Reduction related to BMU Functions
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APPLICATION OF STRAIGHT PROJECT We performed the case study of this asteroid sample return mission based on the STRAIGHT project. This mission belongs to the generation- V planetary mission and requires highly autonomous operations due to the large communication delay and the low communication data rate . Under the severe constraint of spacecraft weight, the information handling should be performed effectively, in the full use of information on spacecraft itself and the asteroid collected by various sensors. BMU with only 9kg weight plays a role in all the information handling in the spacecraft. The power requirement is relatively large (350W at 2 A.U.) because the electrical propulsion is used in the mission. It is very effective to increase the performance of the solar paddles. The light- weight rigid solar paddle with 40W/kg, which is higher by a factor of than conventional one, will be developed to the mission. Also, the refined NiMH battery with 50Whlkg will be used The advancement of communications subsystem include light-weight CFRP parabolic high gain antenna, the light-weight S- band transponder, and the X-band power amplifier with 30% conversion efficiency. Weight and power budget for the asteroid sample return mission is shown in table 1 on the basis of above-mentioned technologies. The reasonable system margin (64.9kg,21.6%) is shown on the table 1. All these advanced technologies make the mission feasible with enough margin of resources. CONCLUSIONS ISAS will perform several planetary missions with our M- V vehicle which is currently being developed. These planetary missions evolve from relatively simple flyby missions to more sophisticated missions such as the asteroid sample return, which this paper deals with. In order to perform these advanced planetary missions under severe constraint of spacecraft weight, the advanced technologies should be exploited to develop light- weight spacecraft with high performance. We have initiated the "STRAIGHT" project for this purpose. The spacecraft function would be devided into four functions; information handling, energy handling, area -dependent function (aperture antenna or sensor), mass-dependent mechanical function. Then we proposed the conceptual flow chart of spacecraft weight reduction. Based on the basic concept, we discussed our present strategy for spacecraft weight reduction as follows. First priority is to reduce weight of the information handling subsystem down to 1/3 or less. In parallel, the other subsystems (power requirement, energy handling, antenna/sensor, and structure/propulsion) do their best to reduce the weight or to improve the performance by a factor of 20% or more. These advance of spacecraft technologies would lead to reduction of the spacecraft weight in half.We proposed the Bus Management Unit as an integrated information handling instrument. Several key technologies in the other three functions are described in the paper. We performed the case study to apply these advanced technologies to the asteroid sample return mission. It is confirmed that these advanced technologies make the asteroid sample return mission feasible even under severe constraint of spacecraft weight. REFERENCE (1) (2) (3) (4)
K.Hirao and T.Itoh : "Project Overview and Highlights of Suisei and Sakigake", Advances in Space Research, Vo1.5, No.12, 1985, pp.55-64 . I.Nakatani et al. : "Planet - B: Japanese Mars Mission", IAF-93-Q.3.400, IAF, Graz, Austria (1993). H.Mizutani et al. : "Lunar Interior Exploration by Lunar Penetrator Mission", IAF 90-039, IAF, Dresden , GDR (1990). K.Aaron: "Elastic Scaling of Small Structures", 7th Annual AIAA/USU Conference on Small Satellites, Logan, Utah, USA (1993).
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Table 1 Weight and Power Budget Table Subsystem/Component
Weight (kll:)
Power Solar Cell Panel Baterry Shunt Dissipator Communicat ion High Gain Antenna 2-axis Gimbal System S-band Transponder S-band Diplexer S-band Switch X-band Transmitter X-band Power Amp. Low Gain Antenna Data Handling Bus Management Unit Data Recorder Launch Ope ration Attittxle/Orbit Control Moment mil Wheel 2-dim. Sun Sensor Opt ical Nav. Cam era 2-axis Gimbal Syst em Laser Altimeter Star Tracker Rate Gyro Assy. Accelerometer Propulsion Reaction Control System Ret urn Tri p Motor Electric Propultion Stntcture Integration Mechanical Thermal Wireharness Balance Weight Mission Payload Sampler Target Plate Re- entry Capsule Scientific Instrument Ory We4!ht Total Fuel ( N2 H4IXe) System Margin Total Weil!'ht Heater Total Power Generation Power @ 2AU • EP
Electric P ropulsion
40 .0
sr0.0
Power (W) Ep· off on 0.0
35.0 2.2 2.8 15.7
190/0Si 40W/kg NiMH 50WWkg 65 .4
20 .4
q, 1.5m
7.2 4.0 2 0.6 0.1 1.2 0.2 0.4 10.0
15.0 22.8
Tx-Power 5W / Eff.=30%
6.0 0.0
11.0 16.6
Tx-Power 5W / Eff.=30Oh
27.5 1.7
132Mbits
29 .2 27 .5 1.3
0.0 16.7
13.7 76 .9
250.0 19.0 45 .0 25.0
24.6 30.3
22.0 2Nms x 4 3.5 6.0 2 unit (Wide!Te le) 10.0 17.0 6.0 10.0 2 unit 2.4 4 unit
7.2 3.5 0.0 0.0 0.0 6.0 0.0 0.0
8.0 1.5 1.0 1.5 2.0 6.0 0.7 0.4 89.0
Meshed FRP
0.0 14.4
28.8 9.0 1.0
4.0 2 1.1
Remarks
70 .0 0.0
70.0
250.0
0.0
IN x 12Thrulters 382m/s 4.9mN
0.0
0.0 11.5 8.8 7.0 3.0
27 .0
0.0 5.0 1.0 15.0 6.0
TBD TBO TBO
0.0 0.0 0.0 0.0
TBO
259.7 50.0 55.3
N2H4 17Om/slXe 1600m1s
365.0 20 .0 335 .9
20 .0 268.4
350.0
1.4kW @ IAU