Chapter 5:Spacecraft and Mission Design

The spacecraft characteristics as the baseline are summarized in Table 5.1. Key itemsof the characteristics are described in following sections.
Table 5.1. Characteristics of the baseline spacecraft.
Spacecraft layout
WeightSize
Launch vehicle
Orbit
Power
Command & Telemetry

Attitude control

Science operations

Telescopes mounted on the optical bench, which is on the Bus
module. See Figure 7.1.
4930 kg (at liftoff), 2330 kg (dry weight)
3.2 m x 3.2 m x 7.4 m, excluding solar array paddles
JAXA H-IIA (type 202)
An inclined geo-synchronous orbit (Baseline)
A sun-synchronous polar orbit (optional)~3,500 W (maximum)
Uplink and housekeeping downlink: S-band
Downlink: X-band 16-QAM (16Mbps)
Ka-band QPSK (80Mbps, optional)
3-axis body pointing with high accuracy
Image stabilization system in some telescopes
A function to change the pointing around Z-axis for matching
spectrograph slit direction to the observing target.
Hinode operation scheme, with a final target selection around 10
to 15 minutes before the uplink in cases where this is essential forthe science.

Two candidates exist for the X-ray imaging telescope. Here we assume thephoton-counting grazing-incidence telescope, because it will require larger size andmass and higher telemetry rate.

5.1. Spacecraft Layout
The layout concept of the Solar-C Plan-B spacecraft is shown in Figure 5.1.
Technical heritage acquired with the development of the Hinode (Solar-B) spacecraft isused as much as possible, including the design concept of the spacecraft structural
layout. The three telescopes are mounted with mounting legs on the spacecraft OBU(Optical Bench Unit), which is a CFRP-made cylindrical structure providing stablecondition to the telescopes. The telescope portion of the UV-Visible-IR telescope(SUVIT) is equipped inside the OBU cylinder. Its focal plane instruments and the othertwo telescopes are mounted on the external wall of the OBU. The mounting legs are
made from low-thermal expansion materials, and provide 6-axis freedom kinematic
mount to the telescopes. The OBU is mounted on the spacecraft bus module with a truss
structure in order to isolate the OBU and telescopes from the influence of the thermal
deformation of the bus module.
It is important to minimize co-alignment changes among the three telescopes andthe attitude sensors against the launch load and the in-flight thermal environment. The
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three telescopes and the attitude sensors (UFSS and IRU) are mounted on the samestructural component, i.e., OBU, to achieve precise co-alignment among them.
The size and mass of the telescopes used for the initial design are listed in Table 5.2.
The values in this table are the initial guess based on the aperture size and focal length
currently under consideration for each telescope. The values contain significant
uncertainty and will be revised as the design is in progress.

Figure 5.1. SOLAR-C plan-B spacecraft
Table 5.2. Size and weight of the telescopes used as the initial parameters.
Instruments
Telescope
UV-Vis-IR
Spectrographtelescope
ImagerUV/EUV spectroscopic telescope
X-ray telescope (photon counting)
Total

Size (mm)ϕ2300 x 4300
1000 x 400 x 3200
1000 x 400 x 3200
1000 x 500 x 5000
400 x 400 x 4500

Weight (kg)
500
100
100
150
150
1000

The initial mass estimate result is shown in Table 5.3. The total mass of all thetelescopes (1,000 kg) is a very rough value (upper-side value) and this will be revised as
the telescope design is in progress. The spacecraft bus system is about 1,260 kg. The
propellant for moving the spacecraft from a geo-transfer orbit (GTO) into an inclined
geo-synchronous orbit (GEO) and maintaining the GEO orbit is counted as about 2,300
kg, which is the initial guess based on the information available on the website ofMichibiki (QZSS), which has an inclined geo-synchronous orbit (GEO). The reasons
why we are considering a GEO orbit as the primary option can be found later.
JAXA H-IIA 202 rocket is assumed as the baseline launch vehicle. This rocket iscapable of installing a 4-ton spacecraft into a GTO orbit. The initial mass estimateshows about 500 kg over the 4-ton, and we will need to make efforts to reduce the mass
by looking at the accuracy of mass estimates. Note that H-IIA 204 rocket, with foursolid rocket boosters, can install a 6-ton spacecraft into a GTO orbit, although it is much
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more expensive.
Figure 5.2 shows the dimension of the spacecraft, which can be fitted in the
Model-4S fairing envelopes. The height is about 7.4m. The front-end position of the
X-ray and UV/EUV telescopes may be located much below the front position of the
UV-Visible-IR telescope, if the 4S fairing is used. This may give a contamination risk to
the UV/EUV and X-ray telescopes, because these telescopes entrance views the outerside surface of the UV-Visible-IR telescope. In the design phase, we should carefully
evaluate this layout and contamination issue in order to minimize the contamination
risk.
Table 5.3. Initial mass estimate result.
Components
Mission instruments
Structures
Bus module
Optical Bench
Thermal control system
Data handling (space wire) system
S and X-band
Communication
Ka-band (optional)
Electric power system
Solar array paddles

Weight (kg)
1000
340
70
100
10
21
TBD
142
120

Attitude & Orbit control system

100

Chemical propulsion system

170

Other peripheral hardware, such as
electrical cables and mechanical parts
Total (dry weight)
Propellant

Remarks
Table 7.2
Based on a standard Bus module
Sizing from Hinode optical bench
Based on Astro-H (space wire) system

Based on Section 7.2.3.

2 wings, 2 panels for each wing.
Including MIB (momentum wheel
isolation bench, 30kg)
Including apogee kick motor (~500N
thruster, section 7.3)

185
2258
2300

Total (at liftoff)

Based on the information available on theMichibiki (QZSS) website.

4558

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Figure 5.2. The spacecraft layout in the 4S fairing of the JAXA H-IIA rocket

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5.2. Science Telemetry
5.2.1. Data rate estimate
Table 5.4 summarizes the data rates estimated from the three mission instruments.
This is the initial guess based on the preliminary science design including the format
size of detectors and typical exposure cadence. Details of the estimate can be found in
Appendix 7-A. The table tells that high rate science telemetry is required to acquire
spectroscopic and polarimetric data with high cadence and resolution. “Standard” caseis for continuous observations without interruption, producing data with a constant rate.
“Burst” case means that the data is produced with highest data rate, but the duration ofsuch observations is limited.
The minimum requirement is that the data can be acquired constantly with the data
rate (the rate after compression) of about 8 Mbps in total. According to evaluation of
Hinode data, lossy compression is acceptable to most science cases, excepting X-ray
telescope’s photon counting data, although the effect from lossy compression needs tobe evaluated in the future. Thus, we assume to apply lossy compression to the data.
Also, much higher rate needs to be handled on board to have burst caseobservations. Assuming lossy compression can be used for burst case, the data output
with 570 Mbps (before compression) and 143 Mbps (after compression) should be
handled properly on board.
Table 5.4. Estimated data output rates from the mission instruments.
Instruments
Solar UV-Vis-IR telescope
UV/EUV spectroscopic
telescope
X-ray telescope(photon counting)Total

no comp.
14.4
320
7.0
140
6.9
110
28.3
570

standard
burst
standard
burst
standard
burst
standard
burst

data rate (Mbps)
lossless comp.
lossy comp.
6.1
2.5
128
48
3.0
1.5
60
30
3.5
55
12.6
7.5
243
143

5.2.2. Data handling on board and data recorder
The overall flow of the science data on board is described in Figure 5.3. It is thebaseline that data compression is performed in each telescope before producing the data
into the spacecraft bus. Since different types of data products (imaging, spectroscopic,
and photon counting) are generated in the three mission instruments, it would be good
to use different compression algorithms more suitable for the properties of data products.
After compressing and packetizing the data, each instrument sends compressed data to
the spacecraft. SpaceWire may be used for the data interface between the telescopes andspacecraft. There is no mission data processor, which is different from Hinode science5-5data handling system.
When we run continuous (i.e. 24 hours every day) observations with the 8Mbps
data generation rate, the total data amount is about 700 Gbits per day. Since themaximum duration of the ground station support at USC may be less than 8 hours, thedata needs to be recorded in the data recorder. The capacity of the data recorder should
be enough to allow us to have continuous observations for 24 hours every day.
Moreover, the data recorder allows us to run the burst observations; Data produced
burstly is stored in the data recorder, followed by downloading the stored data after theacquisition when the spacecraft is linked to a ground station. The capacity of the data
recorder will be determined by considering the observation plans and the capacity
reality.

Figure 5.3. Overall data flow estimate

5.2.3. Science Telemetry System
The high-speed telemetry channel is absolutely required for the Solar-C Plan B
mission to transfer the huge amount of data expected from the onboard telescopes. The
requirement is to design the telemetry and data storage system to enable downloading
the volume of the data produced constantly with the rate of about 8 Mbps.
An initial trade-off study suggests either an X-band telemetry system (16Mbps,
16-QAM modulation, with semi-directional low-gain antennas), or Ka-band telemetry
system (80[TBD]Mbps, QPSK modulation, with directional high-gain parabolaantenna), or the both of them. Each telemetry system is briefly described below.

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X-band telemetry system (Figure 5.4)
X-band telemetry is easy to handle, but the problem is that the allowable bandwidthis restricted within 10MHz for near-earth space science applications in X-band
frequency band, according to a recommendation by the ITU’s SFCG (Space Frequency
Coordination Group). If we use the QPSK (phase shift keying) modulation, same as
Hinode’s X-band modulator, the telemetry speed needs to be below about 8 Mbps tomeet the signal bandwidth restriction. Instead, the 16-QAM (quadrature amplitude
modulation), which is widely used on the ground communication and will be tested in
one of Japanese technology demonstration satellites, can increase the telemetry speed to
16 Mbps.
If we have 12 hours for the telemetry downlink in total per day, the telemetry system
can transfer all the data acquired with 8Mbps continuously for 24 hours. We assume
USC 34m antenna as the primary antenna, and the system can be designed to have apositive link margin. Considering other spacecraft operations at USC 34m, the durationof downlink may be restricted below 8 hours. Thus it would be helpful to haveadditional downlinks on NASA and ESA stations, if antennas are available in Asia,
Oceania, and Pacific Ocean area. It is also noted that a huge mass memory (recorder) isneeded to record the data taken during non-contact period.

Figure 5.4. X-band telemetry system
Ka-band telemetry system
Another possibility is to newly explore the usage of Ka-band in Japan. Higher
telemetry rate is possible because a wider bandwidth can be utilized in the Ka band by aspacecraft. Two candidate bands are identified; 1) ~17GHz, a portion of groundequipments available at USC can be used to minimize the cost. 2) ~26GHz, this band isused at the Near Earth Network (NEN) of NASA. It is needed to newly prepare thesystem for receiving signals at this band in Japan, if the telemetry is received at Japan.
When the bandwidth is 100 MHz, 80 Mbps is achieved with QPSK. The downlink
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duration can be reduced to a few hours, although a more huge mass memory (recorder)
is needed.
A concern is that attenuation of signals due to bad weather (rains and heavy clouds)
in Japan affects quality of the data and gives complication to real-time daily operations.
The both X-band and Ka-band systems work in case of the GEO orbit; X-band
16-QAM (16Mbps) as a baseline and Ka-band QPSK (80Mbps) as an option, though wehave to implement the Ka-band capability at a ground station. However, it is impossible
to adapt the X-band system if the spacecraft is in a sun-synchronous polar orbit, because
about 15 passes are too small to download all the data. The only way is to adapt theKa-band system, but a Ka-band antenna is required in the polar region, such as at
Svalbard. At least 144 minutes are needed as the total contact duration to download all
the data with 80Mbps and this duration corresponds to about 12 station contacts. During
each contact (duration ~ 12 minutes), the Ka-band high-gain antenna is quickly moved
to direct to the ground station, resulting in a concern of micro-vibration from theantenna mount mechanism. More details of the downlink system for the GEO orbit aresummarized below:
1.

2.

GEO/X-band (baseline) The 16-QAM modulator, which is under development for the ASNARO
mission The bit rate is 16 Mbps, which occupies the band width ~10MHz
 Semi-directional low-gain antenna (angle ~ 60°, 0dbi), which is used in
HAYABUSA. At least 6 (TBD) antennas are mounted to cover any direction,
one of which is selected by switch.
 50W TWTA amplifier, which is under development for the ASTRO-G mission.
 The link margin is XXX dB for USC 34m.
 It is necessary to implement the X-band 16-QAM demodulator at the USC.
GEO/Ka-band (optional) The QPSK modulator and the bit rate is 80 Mbps (band width=100MHz)
 ϕ 0.3m high-gain antenna (32dbi) + 0.5W SSPA, whose pointing is driven by a
two-axis mount mechanism.
 The link margin is XXX dB for USC 34m.
 It is necessary to make the Ka-band (17GHz) at the USC available. Anotherpossibility is to make use of Ka-band (26GHz) with NASA or ESA groundstation.

5.3. Orbit
As briefly discussed in section 5.2.3, two orbit candidates are considered for thePlan B spacecraft. Currently, an inclined geo-synchronous orbit (GEO) is considered as
the baseline, with a sun-synchronous polar orbit (SSO) as the backup.
Geo-synchronous orbit (GEO): Same as Solar Dynamic Observatory (NASA)
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 Altitude: 36000km, inclination: <30°, period: 1 day
Sun-synchronous polar orbit (SSO):
 Same as Hinode, TRACE, and IRIS.
 Altitude: 680-800 km, inclination: 97-98°, period: 98 min
The GEO orbit would provide much more advantages than the SSO orbit,
although a demerit on the weight of the spacecraft. Advantages are listed below:
 X-band telemetry can be utilized to download the data volume required forobservations, though the duration of contacts is fairly long (8 hours or longer perday).
 Thermal environment is colder in GEO than in SSO because of lower influence ofinfrared radiation from the earth, which much helps the thermal design of the
optical telescope (see section 4.3.4).
 GEO is thermally very stable, providing less orbital variation of thermal
environment. This would reduce structural deformation, resulting in less orbital
variation in pointing stability (section 5.4.1) and temperature fluctuation oftelescope optics.
 A new type of methodology for science operations (section 5.7) can be adapted,
which would much help the narrow field-of-view spectroscopic telescopes toperform optimized observations
 Eclipse seasons are shorter in GEO than in SSO. SSO has eclipse season with
about 20 mins (at maximum) duration every 98 mins orbital period and the season
continues three months every year. The GEO’s eclipse is about 72 mins (atmaximum) duration per day and it continues only one month, twice every year (seeFigure 5.5). A potential demerit is to need a large capacity of the battery forkeeping some components warm without solar heat input.

Figure 5.5. Eclipse duration in the case of GEO. The maximum duration is 72 mins perday. The eclipse season happens twice every year around the equinox. The duration of

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the eclipse season depends on the inclination angle of the orbit.

Figure 5.5: Visibility of the spacecraft from USC that is located at latitude of 30°N
when the spacecraft is in GEO with inclinations.
It is not necessary to install the spacecraft into the geostationary orbit (GSO) in which
the spacecraft is located directly above the equator where the inclination angle is 0°.
When the orbit is at non-zero inclination, the duration of the eclipse season can be madeshorter than one month (Figure 5.5). The spacecraft is no longer stationary when viewed
from a ground station at the non-zero inclination. The spacecraft is always visible at anelevation angle higher than 45° at USC (latitude 31°N) when the inclination angle isless than 30° (Figure 5.6), which is high enough for continuous contact without
interruption.
The most critical issue of GEO is that an apogee kick motor is required to install the
spacecraft from GTO to GEO. The engine and its propellant take half of the total weightat the lift-off as is shown in Table 5.3.

5.4. Pointing Stability and Attitude Control
5.4.1. Requirements on the pointing stability
We set targets of angular resolution better than 0.1 arcsec in the Solar
Ultraviolet-Visible-IR telescope, and better than 0.2 - 0.5 arcsec in the UV/EUV and
X-ray telescopes. The angular resolution requires high stability of the body pointing,otherwise images are blurred due to residual pointing jitter. The requirements on the

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pointing stability are thus derived based on the requirements on the angular resolution
and exposure duration to obtain a data set (an image or set of images), and aresummarized in Table 5.5. In Figure 5.7, the requirements on the pointing stability forboth HINODE and SOLAR-C are indicated and compared as a function of frequency.
Table 5.5. Requirements on the pointing stability of the mission instruments.
Instruments

Time scale

Solar UV-Vis-IR telescope (*)

1 sec
10 sec1 hourMission life
0.5 sec
5 sec
1 hourMission life
1 sec
10 sec1 sec
1 min
1 hourMission life

UV/EUV spectroscopic
telescope

X-ray
telescope

Normal incidence(*)

Grazing incidence

Instruments

Time scale

Requirements(θx/θy)
0.015
0.015
2
20
0.3
0.3
2
32
0.1
0.1
0.3
0.7
8
32
Requirements
(θz)
50

Unit
arcsec 3σ
arcsec 3σ
arcsec 0-p
arcsec 0-p
arcsec 3σ
arcsec 3σ
arcsec 0-p
arcsec 0-p
arcsec 3σ
arcsec 3σ
arcsec 3σ
arcsec 3σ
arcsec 0-p
arcsec 0-p

Unit

Solar UV-Vis-IR Telescope
1 hourarcsec 0-p
UV/EUV spectroscopic
1 hour100
arcsec 0-p
telescope
Normal incidence
1 hour50
arcsec 0-p
X-ray
telescope
Grazing incidence
1 hour100
arcsec 0-p(*) The pointing stabilities are to be achieved with the usage of an image stabilization system inside
the telescopes.

It is important to identify critical items to achieve the pointing stability
requirements by comparing on-orbit pointing performance of HINODE that achieved
the highest pointing stability in Japanese space missions. We evaluate the pointing
stability of HINODE in the frequency domain using telemetry data taken with the Ultra
Fine Sun Sensor (UFSS) and the mechanical gyroscope (Inertia Reference Unit, IRU)
that are components of the attitude and orbit control system (AOCS) on-board HINODE.
They provide residual pointing errors with a temporal sampling of 0.5 sec, which
provides the pointing stability in the frequency domain lower than 1 Hz. The correlation
tracker (CT) on-board HINODE provides signals of pointing errors with a temporal
sampling of 580Hz, which is used to evaluate high frequency disturbances in thefrequency range between 0.1 Hz and 290 Hz. The on-orbit pointing stability evaluatedwith UFSS, IRU, and CT of HINODE is shown in Figure 5.7, as comparison with therequirements on the pointing stability for HINODE (top) and SOLAR-C (bottom). It isfound that the AOCS on-board HINODE delivers better performance than therequirements in the frequency range lower than 0.1 Hz (upper limit of the AOCS control
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band). It is important to note that, without the image stabilization system, it is difficult
to realize the good spatial resolution with SOT because the pointing stability of about
0.1 arcsec (0-p) at 0.1 Hz achieved with the AOCS is worse than the requirements of
SOT. The image stabilization system has a crucial role in the reduction of the attitude
jitter in the frequency range between 10-2 Hz and 10 Hz.

Figure 5.6. On-orbit performance of the pointing stability in HINODE in comparison
with the requirements of the pointing stability for the mission instruments aboard
HINODE (top) and SOLAR-C (bottom). The thick, dashed and dotted lines represent
the requirements as a function of frequencies while the lower boundaries of the shaded
belts indicate conditionally acceptable level, which is defined as one third of therequirements. The blue filled circles show the pointing stability evaluated with UFSS,the red open circles are that with IRU. The open cyan squares indicate the pointing
stability evaluated with CT in the case of servo-on, and filled orange squares are that in
the case of servo-off.

By comparing the on-orbit pointing performance of HINODE with therequirements for SOLAR-C, we have identified three critical items for further
improvements necessary in the Plan B spacecraft:

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(1) Orbital variation (10-3 – 10-5 Hz)
It is caused by misalignment between the attitude sensors and the telescopes due tothermal deformation of structures supporting them. In the case of SSO, temperaturefluctuation is mainly induced by significant influence of infrared radiation from theearth. In SOLAR-C, the orbital variation is to be reduced significantly thanks to thelower temperature fluctuation at GEO.
(2) Degradation of the pointing stability at around 0.1 Hz
There is an unexpected degradation at around 0.1Hz in the on-orbit pointing
performance of HINODE. The frequency of 0.1 Hz corresponds to the upper limit ofcontrollable frequency of the AOCS. The stability is 0.1 arcsec (0-p), which is notenough to achieve the spatial resolution better than about 0.3 arcsec. It is necessary
to clarify the cause and to improve the AOCS not to induce the degradation in
SOLAR-C. If it is difficult to improve the pointing stability at around 0.1Hz onlywith the AOCS, the mission instruments have to be equipped with an image
stabilization system inside them.
(3) High-frequency disturbances (>100 Hz)
The disturbances at the frequencies higher than 100 Hz have to be suppressed toachieve the resolution of 0.1 arcsec in SUVIT of SOLAR-C. They are induced bystructural resonance between the telescope structures and disturbances of momentum
wheels (MWs) and mechanical gyros. Because the image stabilization system aboard
HINODE is not effective to such high-frequency (>20 Hz) disturbances, one possible
approach is to improve the image stabilization system to have a broader response in
the frequency as high as 100-200 Hz. The other approach is to suppress generation
or transmission of the disturbances from MWs and the mechanical gyroscopes,
which is technically possible if MWs are mounted on a vibration isolation bench.

5.4.2. Attitude control system
The attitude control system provides three-axis stabilization of the spacecraft to point
to an observing target on the solar surface. We can basically use the same configuration
as HINODE, which achieved the highest pointing stability in Japanese space missions,but some modification is necessary in order to improve the items required for further
improvements of the pointing stability described in section 5.4.1. Major components inthe attitude control system are listed below.


Attitude and orbital control processor (AOCP)
 The attitude computer is based on Hinode and other JAXA spacecrafts.
Ultra Fine Sun Sensor (UFSS) The sensor precisely measures the direction of the Sun with sub arc second
accuracy.
 The performance of UFSS on-board HINDOE is enough for the Solar-C
requirements.

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









Fine Sun Sensor (FSS) The planned sensor has been used on Planet-C and so on.
Star Tracker (STT) The start tracker measures the roll around the direction of the Sun.
 The performance of the star trackers used in other spacecrafts meets the
Solar-C requirements.
Tuned Dry Gyroscope (TDG) The gyroscope measures the angle change of the spacecraft attitude as aninertia reference unit (IRU).
 TDG has been used in many space missions in Japan. We have to investigate
the possibility to reduce the disturbances caused by the mechanical gyroscope.
 As an alternate option, fiber optic gyroscope (FOG) is considered for a
disturbance-free gyroscope.
Geomagnetic Aspect Sensor (GAS) The sensor may be purchased from an aerospace company.
Momentum wheels (MW) Momentum wheels are used as actuators of the spacecraft attitude.
 The baseline for momentum wheels is same as used in Hinode.
 For suppressing transmission of the micro-vibration disturbances, vibration
isolation bench may be required to mount momentum wheels.
Magnetic torquer (MTQ) The sensor may be purchased from an aerospace company (following theprevious missions).

5.4.3. Flexibility on the role attitude
In the Hinode case, the spacecraft role is always fixed to align the spacecraft Y-axis,
i.e., Y direction of CCD detectors, to the solar North-South direction. For the Solar-C
Plan B mission, it would be useful to allow changing the spacecraft role attitude,
depending on the nature of observing target. For example, when the slit of spectroscopic
telescopes is placed along a coronal loop structure, the spectroscopic telescopes canrecord spectral lines as a function of spatial position along the loop structure with very
limited number of slit scan positions (or in sit-and-stare mode). This can provide highcadence of observations, allowing us to track energy flows along magnetic fields.
This capability will be considered in the AOCS functional design. It would be notdifficult in implementing this capability from viewpoint of AOCS. However, thiscapability would give complication to the command uplink and telemetry downlink
operations. We may need more number of onboard antennas to cover any directions forthe communication link with a ground station. Or, insufficient communication link levelmay be given depending on the roll angle.
5.5. Spacecraft thermal design to dump heat load from 1.5m telescope
In the thermal design of the spacecraft system, we basically follow the same concept
as adapted in HINODE, which is summarized in the following:
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



The OBU is made of low-thermal expansion and high-thermally conductive CFRP.
The telescopes are thermally isolated from OBU and the spacecraft bus.
The OBU is thermally isolated from the spacecraft bus.

However, as described in section 4.3.4, some part of the large heat load given to
the primary mirror (M1) of the UV-Visible-IR telescope (SUVIT) needs to betransferred toward the spacecraft bus direction (-Z direction). With having the heat path
toward the bus direction, the temperature of M1 can be kept well within the acceptabletemperature range. Thus, a thermal interface is required between the SUVIT’s primary
mirror and the spacecraft. A preliminary thermal concept is to have heat pipes to transfer
the heat to the radiators equipped at the bus side (Figure 5.8). The heat load at M1 isdumped to a cold plate located behind M1 with radiation. Heat pipes connect the coldplate to the radiators on the side of the bus or OBU structure. The heat load of about 100– 200 W is required to transfer via heat pipes. A preliminary study shows that the size ofthe radiators necessary for 100-200W heat dump can be accommodated either on theside of the bus module or OBU structure.

Figure 5.7. Thermal design concept to dump the heat load from M1 of the SUVIT-OTA.
5.6. Spacecraft power system
The spacecraft has solar array panels that generate the electric power. Considering
the size of the Bus module and telescopes, 4 solar array panels would be a baseline.
They are attached on the Bus; either two wings, each of which has two panels, as shownin Figure 5.1, or 4 panels (wings) at each face of the Bus module.
The 4 panels can generate the power up to ~3500W. The power budget will be
evaluated in the design phase, but we will have no critical concerns on the power budget,from our former experience on Hinode (~1100W).
The proposed orbit provides a long continuous duration of non-eclipse condition.
For eclipse periods and initial launch operation, the power system needs a power battery
to provide the electrical power to the onboard components. The size of buttery will be
determined after the initial evaluation of power budget in the design phase.

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The power system provides an unregulated 50V power to all the components onboard; 50-52 V in non-eclipse period and 30-49V in the eclipse period, which is same asused in Hinode.
5.7. Science operations
Since the spacecraft is viewed for a long duration from a ground station (USC),
science daily operations may become more flexible than Hinode’s operation. The
following methodology may be added to the science planning scheme; an observer inthe operation room checks images in real time, and make commands for adjusting thetelescope pointing to the small area of interest, and then the operator uplinks thecommands to the spacecraft during a contact (Figure 5.9). This kind of operations would
be very powerful, because telescopes may have the field of view smaller than Hinodeand the slit position for spectroscopic telescopes is very critical to capture thephenomena of interest.

Figure 5.8. A new methodology of science operations that can be considered in the
GEO case
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Appendix 5-A. Data rate estimate
We here estimate data production rates from the mission instruments using their preliminary
design including formats of their detectors and typical exposure duration in order to clarify how
much telemetry rate is required to download the data from the spacecraft to the ground. We define
two typical observing cases for each instrument; one is for continuous observations without
interruption (“standard case”), which produces a lower data rate, and the other is for short-term burst
observations (“burst case”), which produces a higher data rate. In the estimation, we assume a bit
depth of their data products, and how much the data amount can be shrunk with lossless and lossy
compression algorithms. Details are summarized in the following.


Solar UV-Vis-IR telescope (SUVIT) / Spectrograph package
 Standard case:- In this case, we can achieve high signal-to-noise ratio of about 104 with deep
integration for measuring chromospheric magnetic fields.
- Number of pixels of an image set:1024 (spatial) x 512 (wavelength) x 4 (Stokes) x 2 (orthogonal pol.)= 4 Mpixel- Bit depth:
16bit/pixel (no comp.), 7bit/pix (lossless comp.), 3bit/pix (lossy comp.)- Temporal cadence:
10 sec at each slit position
 6.4 Mbps (no comp.), 2.8 Mbps (lossless comp.), 1.2 Mbps (lossy comp.)
 Burst case:- In this case, we can achieve standard signal-to-noise ratio of about 2x103 with 1 sec
integration for rapid slit scanning.
- Number of pixels of an image set:
2048 (spatial) x 1024 (wavelength) x 4 (Stokes) x 2 (orthogonal pol.)= 8Mpixel- Bit depth:
16bit/pixel (no comp.), 6bit/pix (lossless comp.), 2bit/pix (lossy comp.)- Temporal cadence:
1 sec at each slit position
 128.0 Mbps (no comp.), 48.0 Mbps (lossless comp.), 16.0 Mbps (lossy comp.)

Solar UV-Vis-IR telescope (SUVIT) / Imager package
 Standard case:- In this case, we get minimum wavelength sampling for Dopplergrams with thenarrow-band filtergraph with reduced spatial and temporal sampling.
- Number of pixels of an image set:
2048 x 2048 (spatial) x 5 (wavelength)= 20 Mpixel

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- Bit depth:
12bit/pixel (no comp.), 5bit/pix (lossless comp.), 2bit/pix (lossy comp.)- Temporal cadence:
30 sec for each wavelength scan
 8.0 Mbps (no comp.), 3.3 Mbps (lossless comp.), 1.3 Mbps (lossy comp.)
 Burst case:- In this case, we get full wavelength sampling of a spectrum line for
imaging-spectroscopic observations with the narrow-band filtergraph with full spatial
and temporal sampling.
- Number of pixels of an image set:
4096 x 4096 (spatial) x 10 (wavelength)= 160 Mpixel- Bit depth:
12bit/pixel (no comp.), 5bit/pix (lossless comp.), 2bit/pix (lossy comp.)- Temporal cadence:
10 sec for each wavelength scan
 192.0 Mbps (no comp.), 80.0 Mbps (lossless comp.), 32.0 Mbps (lossy comp.)


UV/EUV high-throughput spectroscopic telescope Standard case:- In this case, we get a reduced set of spectrum lines with deep integration.
- Number of pixels of an image set:
2048 (spatial) x 128 x 10 (wavelength)= 2.5 Mpixel- Bit depth:
14bit/pixel (no comp.), 6bit/pix (lossless comp.), 3bit/pix (lossy comp.)- Temporal cadence:
5 sec at each slit position
 7.0 Mbps (no comp.), 3.0 Mbps (lossless comp.), 1.5 Mbps (lossy comp.)
 Burst case:- In this case, we get a full set of spectrum lines with shorter integration for rapid slit
scanning.
- Number of pixels of an image set:
2048 (spatial) x 256 x 20 (wavelength)= 10 Mpixel- Bit depth:
14bit/pixel (no comp.), 6bit/pix (lossless comp.), 3bit/pix (lossy comp.)- Temporal cadence:
1 sec at each slit position
 140.0 Mbps (no comp.), 60.0 Mbps (lossless comp.), 30.0 Mbps (lossy comp.)

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

X-ray imaging (spectroscopic) telescope Assumptions:- Perform tagging for each photon-incident event- Tagging information
 Pixel address (X, Y) Photon energy Incident time
- No on-chip/on-board pixel summing. Perform on ground if necessary.
- Data compression may be possible, which reduces data rate to half.
 Standard case:- In this case, we get photon counting imaging spectroscopy in 128 x 128 pixels ROI.
- Pixel address (X, Y):
7 bits x 2 = 14 bits (for 0–127)- Photon energy:
10 bits (0–1023 ch)- Exposure time:
every 1 ms
- Ratio of event pixels:
2 % (i.e., 128^2 x 10% = 327.7 pixels contain events)
 6.9 Mbps (no comp.), 3.5 Mbps (comp.)
 Burst case:- In this case, we get photon counting imaging spectroscopy in 512 x 512 pixels ROI.
- Pixel address (X, Y):
9 bits x 2 = 18 bits (for 0–511)- Photon energy:
10 bits (0–1023 ch)- Exposure time:
every 1 ms
- Ratio of event pixels:
10 % (i.e., 512^2 x 10% = 6553.6 pixels contain events)
 110.0 Mbps (no comp.), 55.0 Mbps (comp.)

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