3 Spacecraft System
The spacecraft system design, including the orbit profiles and operation concept, are presented inthis section. Section 3.1 summarizes the requirements to the spacecraft system derived from thediscussion in the previous chapters. Obviously, the most important requirement of SOLAR-C PlanA is to place the spacecraft at a high solar latitude. Working out the way to reach there dominates
the mission design, which is discussed in Section 3.2. Several options are investigated, and an
option using solar electric propulsion (SEP option) is presented as the primary option in this report.
Based on the system requirements and the orbit design results, the spacecraft system for carryingout the mission is briefly investigated. The objective of the study is to confirm the feasibility of themission mainly from the point of view of the resource budget, and its successful result is given in
Section 3.3. The details of the subsystems are to be shown in Section 3.4, At this stage of interim
report, however, the main focus is on those subsystems that require newly developed technology(communication system, light weight solar panel, and electric propulsion system), and the othersubsystems are only briefly touched.

3.1 Spacecraft System Requirements
The major requirements to the spacecraft system (including orbit design) are,
- To reach the solar latitude of 40° by early 2020’s.
- To support science instruments’ mass of 130 kg (or larger).
- To transmit the scientific data continuously recorded in the rate of 100 kbps to the ground.
The third item is interpreted as the 300 kpbs down link rate under the assumption of 8 hours use of
the ground station per day.
The following requirements are also recognized. However, they are regarded as non-dominating in
the total system design and left for the further study in this report.
- Providing payloads a stable platform for imaging observations- 150 W (TBD) payload power- Regular notification of absolute time for regular-interval exposures required by the helioseismic
instrument.

3.2 Orbit and Mission Profile
The SOLAR-C Plan-A is the first Japanese solar mission that operates in the interplanetary space.
The following new issues in mission profile should be noticed when compared with the otherJapanese (Earth orbiting) solar missions.
- It cruises in an interplanetary orbit, and takes much longer time to reach the final orbit.
- It only carries light payloads even if a heavy launcher (as H-IIA) is used.
- It is in severer radiation environments.
The following new issues from other Japanese interplanetary missions should be noticed as well.
- It requires much higher telemetry rate than ever achieved.

46This item is not only reflected in the communication system design, but also considered in the orbit
design to keep the distance from the Earth to the spacecraft short.

3.2.1 Trajectory Options for SOLAR-C Plan-A
The major mission requirement of SOLAR-C Plan-A is to observe the Sun from high latitudes. The
target maximum latitude is tentatively specified as 40° (Figure 21a). To observe the Sun from thehigh latitudes, the space observatory (spacecraft) must be in an orbit largely inclined with respect tothe ecliptic plane. It is not an easy task to inject the spacecraft into the orbit of this type. A roughestimate shows that the velocity increment required to inject the spacecraft into this largely inclined
orbit is approximately 20 km/s (Figure 21b).

v

Figure 21. Schematics of Trajectory Design Challenge in SOLAR-C Plan-A

In order to achieve this severe mission target, possible trajectory sequences are investigatedconsidering the application of various trajectory manipulation techniques. The items considered are,the geometric relation (i.e. the tilt of the solar equatorial plane to the ecliptic plane), launcher
capacity, planetary gravity assists, and the usage of a highly efficient propulsion system. As a result,four major trajectory options are listed up to achieve the mission objective (Figure 22).
The first option is called “SEP option”, which is characterized by the usage of the solar electric
propulsion (SEP) combined with the Earth gravity assists (EGAs). In this option, SEP is used toincrease the relative velocity (v) at the Earth encounter and the direction of v is changed by EGA
to contribute to inclination increase. As a result of the sequential use of this process, latitude 40°
can be achieved in five years from the launch. The other options are characterized by the usage ofthe planetary gravity assists only, without using SEP. The second option is called “Jupiter option”.
In this option the operation sequence begins by using the Jupiter gravity assist (JGA) to highly
incline the orbit plane from the ecliptic plane. What is unique to this option is the usage of EGAs inorder to reduce the orbit period after JGA, which enables the frequent observation of the solar polar
region. The remaining two options are characterized by the usage of EGAs and the Venus gravity
assists (VGAs). In both options the operation sequence begins with the sequence of EGAs/VGAs to
increase v to the planets. Then, the sequential planetary gravity assists are used to change thedirection of v so as to contribute to incline the orbit. “Venus-1 Option” uses the Venus for theplane change gravity assists, whereas “Venus-2 Option” uses the Earth for this purpose.
47

Among the four options, the first “SEP option” has a major advantage in complying with the mostof the mission requirements (the detail shown in 3.1.2). The option has frequent launch windows(every half year), which should be also counted as a merit from the programmatic point of view. On
the other hand, it must be noted that this option requires an advanced technology, which is not
required in the other options. This point will be discussed in 3.3. In view of the major advantage
mentioned above, in the following part of Section 3, the case of SEP option is focused on.
Alternative Jupiter option with the same final orbit is shown in Appendix B.

Figure 22. Trajectory Options for SOLAR-C Plan-A

3.2.2 Orbit and Mission Profile of SEP Option
Before showing the baseline sequence of the SEP option, a couple of its variants are introduced
briefly. The first method, which is called “Direct Inclining Method (DIM)”, is a simple one whichuses SEP directly to increase the inclination. The second method uses SEP combined with EGA,
which is called Electric Propulsion Delta-V Earth Gravity Assist (EDVEGA). The method uses48

EDVEGA repetitively, and is called “Sequential EDVEGA Method (SEM)”. Both methods were
investigated quantitatively, and it was concluded that, DIM is infeasible from the points of view ofthe spacecraft’s mass budget and the operation time of the ion engine system (IES), whereas SEM
is feasible in these aspects. A slight variation of SEM is also investigated, which applies anadditional Venus gravity assist (VGA) prior to the sequential EDVEGA. This method has anadvantage in reducing the launch energy drastically compared to the original SEM. However, theuse of Venus/Earth gravity assists limits the launch opportunity, which makes it difficult to takeadvantage of the geometrical relation (i.e. the tilt of the solar equatorial plane to the ecliptic plane).
From this point, the usage of VGA is regarded only as a backup option.
As a result of the discussion above, SEM is adopted and used to construct the baseline sequence of
the SEP option. The basic procedure of SEM is described as follows.
1. The spacecraft is injected into the Earth synchronous orbit to re-encounter the Earth after one
year cruise.
2. During the cruise, SEP is used to maximize the spacecraft’s relative velocity to the Earth (v) at
the next Earth encounter. Note that the thrust does not necessarily increase the inclination by
itself. To enhance the efficiency to increase v, an elliptic orbit is used for the cruise orbit.
3. By EGA, the direction of v is changed to contribute to the inclination increase.
4. By the repetitive use of the steps 2 and 3, the inclination is increased step by step.
Before presenting the constructed trajectory sequence, major assumptions used in the trajectory
design should be noted.
The first assumption is related to the launch condition. The assumed launcher is the Japanese H-II
A heavy launch vehicle equipped with a solid motor upper stage. The initial mass of the spacecraft
is assumed to be 1200 kg. Assuming the practical settings of the launch site and the launch
direction, the launcher is capable of injecting the spacecraft into the escape orbit with the excessive
velocity (v) of 7.3 km/s. The launch date is selected to take advantage of the not negligible tilt
(7°.25) of the solar equatorial plane to the ecliptic plane, and they are June 7 or December 8. In thefollowing discussions, June 7 is used as the launch date.
The second assumption is related to the ion engine system (IES). The specific impulse (Isp) of IES
is assumed to be 3800 s, and the maximum thrust of IES is assumed to be 120mN in total. In the
trajectory design, the actual thrust available for the maneuver is constrained by the available power,
which is assumed to decrease as the inverse square of the distance from the Sun.

49

Figure 23. Trajectory profile of SEP option

The third assumption is related to the eccentricity (e) of the cruise orbit. To efficiently increase v,the eccentricity e should be large. However, considering the difficulties in the thermal design of thespacecraft, the eccentricity e is constrained to be less than 0.3 in the trajectory design.
Figure 23 shows the trajectory profile of the SEP option. In Figure 23a, the trajectory of the first
revolution is projected on the ecliptic plane. The figure shows that the spacecraft’s orbit has
eccentricity, and intersects with the ecliptic approximately at the Earth’s position at the launch. The
acceleration vectors indicate that the thrust is used to decelerate the spacecraft at aphelion and toaccelerate the spacecraft at perihelion, which result in the increase of v at the next Earth encounter.
Figure 23b shows the trajectory profile through the sequence. The trajectory is projected on theplane perpendicular to the ascending node direction so that the gradual change of the orbit plane
can be easily seen. It is observed that the first four cycles have asymmetry resulted from the orbit
eccentricity. However, the orbit is finally circularized by EGAs, which results in the symmetry
observed in the final orbit (the orbit whose inclination is the highest).
Table 7. Sequence of events of SEP option

50

The sequence of events of the SEP option is summarized in Table 7. In the table, v denotes therelative velocity to the Earth at EGA, and iSEQ denotes the inclination to the solar equatorial plane.
The spacecraft reaches iSEQ of 30° after the 3rd EGA (3 years from the launch), and finally reaches
iSEQ of 40° after the 5th EGA (5 years from the launch).
Figure 24 shows the profiles of some important parameters of the mission. The top chart indicatesthe points of events and basic operation concept. The following three charts respectively show the
profile of the spacecraft’s distance from the Sun and the Earth, its instantaneous solar latitude, andthe expected down link rate of the scientific data.

cFigure 24. Mission profile of SEP option

In the first three years from the launch, before the spacecraft reaches iSEQ of 30°, the spacecraft
operation is devoted to increase the inclination. IES is operating most of the time, and silent
condition for the scientific observation is basically not guaranteed in this phase. Even in this phase,intermittent suspension of IES is planned for accurate orbit determination, which may be used as51

occasions for observations. In the fourth year, after EGA #3, iSEQ exceeds 30°. From this year
onwards, the period during which the spacecraft is at latitudes higher than 30° is allocated as “theobservation phase (orange area in Figure 24)”, and the IES operation is intentionally suspended.
Even after EGA #3, IES operation is continued while the spacecraft is in the lower solar latitude.
This IES operation and the following two EGAs contribute to the further increase of the inclinationand the circularization of the orbit. Finally, as a result of EGA #5, the spacecraft is injected in to thefinal observing orbit, the circular one-year orbit with iSEQ of 40°.

3.3 Spacecraft System Design
The SOLAR-C Plan-A is the first Japanese solar mission that operates in the interplanetary space.
The following new issues in spacecraft system design should be noticed when compared with theother Japanese (Earth orbiting) solar missions (a part of issues results from the mission design).
- It requires a much longer mission life.
- It is in severer radiation environments.
- It carries light payloads only.
- It requires large solar array panels- It requires fuel to keep three-axis attitude control.
The following new issue from other Japanese interplanetary missions should be noticed as well.
- It requires much higher telemetry rate than ever achieved.
This item is already considered in the orbit design, and is also reflected on the communication
system design. In addition, it should be noted that, as a result of using a liquid fuel rocket (H-IIA),
the mechanical environment at the launch is relaxed from that in case of using solid rocket as
previous solar missions.

3.3.1 Spacecraft Configuration
The mission design provides basic information of the spacecraft configuration. The required thrust
determines the configuration of IES, and the configuration of IES determines the required power
which finally determines the size of solar array panels (SAP). The orbit design provides thegeometrical relation around the spacecraft during the mission, which determines the basic locationof the key components. As a result, the basic configuration the spacecraft is determined as shown in
Figure 25.

52

Figure 25. Spacecraft configuration of SEP option

The spacecraft’s +Z-direction is always pointing to the Sun. The scientific payload assembly ismounted on the +Z panel of the bus module and pointing to the Sun. SAP are deployed from thebus module to the ±Y-direction and fixed to the body, facing the +Z-direction. In the trajectory
design process, IES thrust direction is constrained to be approximately normal to the Sun direction,
which enables IES module to be mounted on the –X panel. The panels not facing to the Sun
direction (±Y, –X, -Z) are used as radiator panels to dissipate the large heat generated by IES
components. The geometrical relation among the spacecraft, the Sun and the Earth varies as afunction of time. Therefore, a high-gain antenna for the high rate data transmission must be
maneuverable, and mounted on the +X-panel.

3.3.2 Mass Budget
In order to confirm the feasibility of the mission, the spacecraft mass budget is estimated. In theinvestigation, the total mass of the spacecraft is assumed to be 1,200 kg (which complies with theassumption of the trajectory design), and the difference with the sum of subsystem masses iscounted as margin.
For most of the subsystems, the composition of the components is considered, and the mass ofcomponents is summed up to the subsystem mass, while the mass of subsystems such as structure,harness, etc. is estimated parametrically based on the previous spacecrafts. The propellant mass for
IES is derived from the trajectory design result.
The resulting mass budget is summarized in Table 8. The positive margin of 106 kg indicates thefeasibility of the mission from the point of the spacecraft’s mass budget.

53

Table 8. Spacecraft mass budget of SEP option

3.3.3 Power Budget
In order to estimate the size and mass of SAP, the required power for the spacecraft is estimated. In
the investigation, two cases of power budget (IES operating / non-operating) are estimated. The
composition of the components is considered, and the powers for the components are summed up to
the subsystem power. The resulting power budget is summarized in Table 9.
Table 9. Spacecraft power budget of SEP option(Heater powers for payload not included.)

The mass of SAP is estimated on the basis of supplying the power for IES operating case at 1 AU.
The use of newly developed light weight SAP is assumed, and the specific value of 100W/kg is54

used in the estimation. The resulting mass of SAP is included in the power supply system in themass budget.

3.3.4 Communication Link Budget
In order to confirm the feasibility of the high rate data link, the communication link budget isestimated under the assumed communication system configuration. In the investigation, two casesof link budget (X-band / Ka-band) are estimated. The maximum distance of 0.56 AU in the final
orbit is used, and NASA/DSN 34m antenna is assumed as the ground station. The resulting linkbudget is summarized in Table 10.
Table 10. Communication link budget of SEP option

Item
Frequency
Transmitter power
Transmit antenna gain
Communication distance
Data rate
Link margin

X-band
8400
40.0
36.5
0.56
300k
1.0

Ka-band
32300
20.0
48.2
0.56
1M
2.0

unit
MHz
W
dBi
AU
bps
dB

Received C/N0
Required C/N0

59.3
58.4

64.6
62.6

dBHz
dBHz

The positive margin of 1 dB for 300 kbps downlink indicates the feasibility of satisfying the system
requirement in the final orbit. It should be noted that this downlink rate is achieved by more reliable
X-band link. In addition, 1 Mbps downlink by Ka-band link is expected with positive margin of 2
dB.

3.3.5 System Thermal Design
In the cruise phase of SEP option, the minimum solar distance (i.e. perihelion distance) is to be 0.7
AU. It is assumed that three μ20 ion engines are simultaneously used even at the perihelion passage.
The amount of heat dissipated from IES is estimated to be approximately 1,800 W in total, and it is
not easy to radiate this amount of heat to space. In order to confirm the feasibility of the thermal
design, system thermal analysis is conducted. In the analysis, the simultaneous use of the three μ20
ion engines at 0.7 AU is assumed as the worst case. The Sun-light is assumed to input from thedirection 5° apart from +Z direction, which causes the heat input from the Sun into the side panelsof spacecraft.
The thermal mathematical model consisting of the spacecraft bus module and ion engine module isconstructed, and a steady state thermal analysis is conducted. The result shows that we can still55

have a feasible solution under this worst condition. The solution requires that concentrated heatdissipation from the ion-engine related components to be averaged at each structure panel by heat
pipes and thermal doublers.

3.4 Spacecraft Subsystems
The spacecraft subsystems are discussed in this section. At this stage of the interim report, theinvestigation gives priority to the subsystems that require new key technologies to achieve thismission. They are Ka/X communication system, ultra-lightweight solar panels, and μ20 ion engine
system. The details of these subsystems are given in Sections 3.4.2, 3.4.3 and 3.4.6 respectively. As
for the other subsystems, the functions required to them are supposed to be achieved with proventechnologies. Their compositions are tentatively assumed based on the previous missions, and they
are only briefly touched in this report.

3.4.1 Data Handling System
The data handling subsystem are supposed to be composed based on the heritage of the previous
interplanetary spacecrafts, Hayabusa or Akatsuki. It is composed of DHU (data handling unit) and
TCIM (telemetry command interface module). DHU manages the automatic/autonomous functionof the spacecraft bus system, collects and records the HK (house keeping) data, and manages DH
network as the host. TCIM interfaces telemetry and command between DHU and thecommunication system. It generates the transfer frame and multiplexes the telemetry packets, andoutput to the transmitter. On the other hand, it modulates the uplink signal input from responder,and output command packets to DHU.

3.4.2 Communication System
Since the spacecraft is in the interplanetary space for the out-of-ecliptic mission, a high-speed
telemetry is mandatory for imaging observations of the Sun. We have tentatively set a straw-man
data recording rate of 100 kbps in X-band considering the continuous helioseismic observations.
This level has been achieved in the NASA STEREO mission in the distance of 0.5 AU from the
Earth. On the other hand, the Japanese interplanetary missions that have so far been flown only
achieved the telemetry rate of less than 10 kbps using 64m-aperture downlink station at Usuda deep
space center. When the downlink duration is limited to 8 hours per day for example, 300 kbps
telemetry rate is required for 100 kbps recording rate. If this should be achieved by a domestic
effort, a development of a high-power transponder will be required in addition to reducing any
transmission losses.
Based on the consideration above, the following design guideline is derived.
- Achieve 300 kpbs downlink rate assuming the downlink duration of 8 hours per day.
- Use X-band system heritage in interplanetary communication for stable and assured downlink.
- Use Ka-band system for a higher-telemetry capability under good atmospheric conditions.
56

- Adopt domestic supplied components which we can have by around year 2012.
- Adopt a high gain antenna of 1.6 m diameter.
- Adopt a high-power (40W) X-band transmitter as a natural extension from the JAXA heritage.
- Reduce feeder loss compared with the previous JAXA interplanetary missions.
- Improve the coding gain by the modification of coding method.
The block diagram of the communication system (with the heritages) is shown in Figure 26.
In order to have a higher data transmission rate or a shorter downlink duration, the use of Ka-band
system is supposed in addition to the X-band system. The downlink of Ka-band telemetry is
possible at the NASA DSN stations to track the STEREO spacecrafts. Since the availability of Kaband is affected by the atmospheric condition on the Earth, the Ka-band alone cannot be a solution
to the requirement set by the continuous helioseismic observations.

KaTWT
KaEPC

X/Ka HGAKaSW-1

HGA Control Unit
X/Ka Feeder

XEPC-1

XTWT-1
HYB

XSW-4
DIP-1
X LGA-1

XEPC-2

XTWT-2

XTRP-1

XSW-3

XSW-2

KaTX-1DIP-2

KaTX-2
XSW-5
XTRP-2XSW-1

XLGA-2

Figure 26. SOLAR-C Communication System

3.4.3 Power Supply System
Power supply system of SOLAR-C Plan-A consists of Power Control Unit (PCU), Series Switching
Regulator (SSR), Solar Array Paddle (SAP), Battery (BAT), and Pyrotechnic Controller (PYC). In
the system, the output voltage of SAP is lowered and stabilized by SSR. The functions of PCU are
SSR control, BAT charge/discharge control and management, and the bus power distribution. BAT
is used for power supply when SAP cannot supply sufficient power. The cases supposed are, thelaunch phase, the attitude anomaly, and the large attitude maneuver. The use of Lithium ion battery
is supposed based on the heritage of the previous interplanetary spacecrafts, Hayabusa or Akatsuki.
SOLAR-C Plan-A spacecraft is propelled by the ion engine system. In order to operate three μ20
ion engines simultaneously, approximately 6 kW of power is required to be supplied by SAP.
When conventional rigid type SAP is adopted, its mass is estimated to exceed 100 kg. On the other57

hand, the reduction of components’ mass is strongly requested as a general nature of interplanetary
spacecraft. From this consideration, the newly developed ultra-lightweight solar panel is used tomeet the requirements of high specific power (W/kg) and low storage volume. This component is
identified as the key technologies to achieve this mission.
The ultra-lightweight solar panel is now under development in ARD/JAXA, and uses newly
developed space solar sheet supported by a frame-type structure (Figure 27). It is targeting thespecific power greater than 100W/kg in panel level. Future perspective of the ultra-light weightsolar panel development is shown in Figure 28. Breadboard model was developed last year, and aseries of development tests (vibration, thermal vacuum, and deployment test) were successfully
conducted. The technical readiness level (TRL) is estimated to be 6+α.

Specific Power, W/kg

Figure 27. Ultra-Lightweight Solar Panel

150
Present state
100W/kg~150W/kg

100>120W/kg

Standard Panel
1m  1m unit

BBM (small sized panel)

50

Panel
Structure

Cell

88W/kg-Framestructure-Framestructure-Lateralpanels-Framestructure-Lateralpanels-Framestructure-Lateralpanels-Pantographic
mechanism

2J

2J

3J

3J

Figure 28. Future Prospects of Ultra-Lightweight Solar Panel Development

As mentioned in Section 3.3.1, in the configuration of this spacecraft, Z-axis (the direction
perpendicular to SAP) is basically pointed toward the Sun. Therefore, SAP is fixed to thespacecraft body, and no driving mechanisms are installed.
Li-ion battery is adopted as a high-density power cell in the Plan-A spacecraft. The overcharge andover-discharge protections are dealt by the Series Switching Regulator (SSR) that has beendeveloped in the JAXA MMO mission. The bus voltage regulation is made by SSR monitoring theelectric current at BAT. The period in which the spacecraft is operated by the power from BAT is58

(1) during initial phase just after launch, (2) during orbit control without using the ion-engine
system, and (3) during the safety state in an attitude control for emergency cases. The eclipse
period in which the Sun cannot be seen does not happen in the SOLAR-C Plan-A orbit except for
the launch operation. The capacity of battery is selected by these conditions and it is found to beabout 27 Ah (TBC) in the first estimate.

3.4.4 Attitude Control System
The requirements for the attitude control are summarized in Table 5. The three-axis stabilization isrequired for imaging observations of an arcsec spatial resolution. The attitude control system that is
tentatively assumed here is what was used in Hinode mission. Since the short-term performance ofthe Hinode attitude control system completely satisfies the SOLAR-C Plan-A requirements, no
improvement is necessary. The external disturbance in the interplanetary space is much smaller andmore stable than that in the low-Earth orbit, so that the main source of disturbance is themechanical motions within the spacecraft. One of the differences between SOLAR-C Plan-A and
the Japanese solar missions in low-Earth orbits is the use of chemical thrusters to unload theangular momentum accumulated in the reaction wheels (RW). According to a simple calculation,
the frequency of the RW unloading is estimated to be once per day in a nominal science operation.
The helioseismic instrument may need to have its own image stabilization tip-tilt mirror tocompensate the long-term pointing change or to stabilize the unexpected low-frequency disturbance.

3.4.5 Chemical Propulsion System
Chemical propulsion subsystem is used for unloading of the RW accumulated angular momentum
and the attitude maneuver which requires large control torque (such as initial attitude establishment
and safe hold mode operation). The thrusters are also used for the short term translational motion
control such as the trajectory correction before the Earth gravity assists. Considering the level ofthe total impulse required, the conventional monopropellant (N2H4) system is assumed at thepresent phase.

3.4.6 Ion Engine System
Owing to the success of Japanese asteroid explorer HAYABUSA, which has an ion-engine μ10
system for interplanetary cruise, ISAS/JAXA has a heritage of ion engines. ISAS is developing thescale-up version of the ion engine thruster, μ20. Figure 29 shows the engineering-model thruster,which is in an endurance running test under a vacuum condition. We assume a SEP system inwhich four sets of ion engines are set on the spacecraft and three out of four can be turned on
simultaneously. In the orbit design we assume the thrust of 40 mN and specific impulse of 3,800 s
for each thruster or the maximum thrust of 120 mN for simultaneous use of three thrusters.

59

Figure 29. Engineering model of μ20 ion engine thruster

The μ20 ion engine system consists of Ion Thruster Control Unit (ITCU), IES Power Unit (IPU),
Propellant Management Unit (PMU), Ion Thruster Assembly (ITA), Microwave Supply Unit
(MSU), and IES Pointing Mechanism (IPM). Xe is the propellant of this engine. Xe ions areproduced in a process of discharge by microwaves and are accelerated by electric fields between
two grids. The role of each unit is summarized as follows:
IPM: This is a pointing mechanism of ITA.
IPU: This produces a high voltage (~2.3 kV) to accelerate Xe ions.
ITA: This unit produces thrust from Xe propellant, microwaves, and high-voltage applied.
ITCU: This unit controls the whole ion-engine system (IES).
MSU: This unit supplies microwaves to ionize Xe for acceleration and to make electrons forneutralization.
PMU: This unit provides Xe propellant to ITA.
Each ion engine thruster produces the thrust of 40 mN. By simultaneous use of three μ20 ion
engines, one of which is currently being developed and tested at ISAS, the maximum thrust of 120
mN is obtained. Some components are set on the spacecraft bus structure, and some are set on the
Ion Engine Module (IEM) that is an independent box located outside the spacecraft bus structure.
IEM has six ion engine thrusters, two of which are redundant engines for unexpected situations.
The IPM connects IEM with the spacecraft, and it also acts as a pointing mechanism that slightly
tilts the whole IEM to control the direction of thrust under combinations of multiple ion engine
thrusters, each has a slightly different thrust and is located at a different distance from the center ofgravity of the spacecraft.
Power Control Unit (PCU) on the spacecraft provides the power of 350 W (1,520 W) with MSU(IPU) for operating each ion engine thruster. 1,870 W electric power is required for operating asingle ion engine thruster. It was 350 W for Hayabusa’s μ10 ion engine system. Larger power
consumption in the SOLAR-C Plan-A system gives large heat generation that needs to be radiatedaway from the spacecraft. The ion engine system in the Plan-A spacecraft produces a huge amountof heat that cannot be radiated away from the surface of IEM alone. In order to have a solution for60

the heat dissipation problem, some components, ITA and MSU, are set in IEM, and othercomponents, ITCU, IPU, and PMU, are mounted in the spacecraft bus structure.
In the following points the system is different from Hayabusa’s ion-engine system. These are all
related to the increase in weight or heat dissipation to achieve high thrust in good propellant
efficiency.
1. Size of ITA from 10 cm diameter in Hayabusa to 20 cm diameter in SOLAR-C Plan-A
2. Operating power for ion acceleration
3. Pointing mechanism at IPM
Currently estimated specifications summarized in Table 11 assume higher specific impulse options
that will soon be applied to the EM thruster under endurance test. Original test conditions of the
EM thruster's ion source were 1,300 V (Screen Voltage), 30 mN (Thrust) and 2,800 sec (Specific
Impulse), respectively, and it accumulated 10,000 hours of operation by Sep. 2010. The technicalreadiness level (TRL) is estimated to be 4-5.

Table 11. SOLAR-C Plan-A Ion Engine System

Propellant
Specific impulse
ThrustScreen Voltage
Number of thrusters
Maximum number of thrusters
that run simultaneously
Size of Ion-Engine Module
Pointing mechanism
WeightWeight of Fuel (Xe)
Power

Xe
3800 s
40 mN/engine
2300 V
4
3
1,2001,500600 mm
Newly proposed linear actuators
Range: 5
162 kg
236 kg5,650 W
Regulated bus:
1,090 W
Non-regulated bus: 4,560 W

Need a development
Estimate in 2010
Estimate in 2010
Estimate in 2010

3.4.7 Structure System
A box-type panel structure is adopted as the spacecraft main body (bus module). The approximate
size of the main body is 2m×1.2m×1.5m. The scientific payload assembly is mounted on the +Z
panel of the bus module, and the rocket attachment ring is installed on the –Z panel. The movable
IES module is installed on the –X panel. Two axes pointing mechanism enables to direct the IES
thrust axis to the spacecraft center of mass. The major deployable appendages are two wings of61solar array panels and Ka/X-band high gain antenna. The solar array panels are installed on the ±Y
panels, and they are latched and fixed to the body after deployment. Two axes gimbaled movablehigh gain antenna is installed on the +X panel, and is deployed after the launch.

3.4.8 Thermal Control System
SOLAR-C Plan-A spacecraft is propelled by the ion engine system. When three μ20 ion engines are
simultaneously operated, the heat dissipated from IES is estimated as approximately 1,800W in
total. The major theme for the thermal control system is to radiate this large amount of heat tospace. In order to achieve this objective, large areas on the structure panels that are not facing with
the Sun (±Y, –X, -Z) are used as radiator panels. Concentrated heat dissipation from IES istransferred and averaged at each structure panel by heat pipes and thermal doublers. The area
outside the radiators on the structure panels is covered with the multi layer insulator (MLI), tointerrupt unnecessary heat input/output.
When not all the ion engines are operated, the reduced heat dissipation from IES is compensated byheaters. The heater control electronics (HCE) automatically controls the duty of the heaters basedon their priority. It enables the appropriate management of the power allocation between thermal
control and IES operation even when the available power is limited around the aphelion.

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