SOLAR

Overview of the Current Baseline of theSolar-C Spacecraft System
Keisuke YOSHIHARA (JAXA)
11 November, 2013
Solar-C Science Meeting
Hida Earth Wisdom Center, Takayama, Japan

Figure 3.8: Physical block diagram of the instrument, showing major components in the instrument and digital

Solar-C Spacecraft System Overview
communication links.

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EUV/FUV High Throughput
Spectroscopic Telescope(EUVST)

Figure 3.9: The instrument opto-mechanical layout

Table 3.6 summarizes unit and assembly key characteristics. Description and design
characteristics of the optics, detectors, structure, and electronics are provided in the sections
below.

Solar UV-Visible-IR Telescope(SUVIT)

Unit/assembly
Telescope unit
Telescope type
Mirror characteristics

Table 3.6: Instrument module design characteristics.
Nominal value/description

Solar-C S/C system

Single-mirror off-axis paraboloid
30 cm diameter, 360 cm focal length, overall < O/75 rms figure, < 5 Å rms
micro-roughness, broadband VUV coating, low-expansion substrate (Zeroduror ULE)

3-22

X-ray Imaging Telescope (XIT)2

Current Baseline Spacecraft SystemSOLAR

Model Specification of Spacecraft SystemSize3.7 m x 3.2 m x 7.1 m
excluding solar array paddles

Weight

4.1 t (at liftoff), 2.3 t (dry weight)

Orbit

Geosynchronous orbit with a non-zero inclination (baseline)

PowerPower generation: 5kW @EOL
Load: 1.5kW (average) + operational heaters

Communication

Mission data downlink: X-band 16-QAM, 24Mbps (baseline)
Ka-band QPSK, 80Mbps (option)
Uplink and housekeeping downlink: S-band

Data rate andvolume

Average mission data rate: 8Mbps
Data recorder storage: 100GB@EOL(3years)

Pointing stabilityand Attitude control

3-axis attitude control with very high Sun pointing accuracy
Image stabilization system in each telescope

Science operationfeatures

Includes close-to-real time access to the satellite and operation
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Solar-C Spacecraft System Overview
UV/EUV spectroscopic telescope(EUVST)

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SAP
Star Tracker (STT)!
Optical bench unit (OBU)!

UFSS tower!

Middle gain antenna(MGA)

Top door!

SUVIT interface unit!

Apogee kick engine!
Heat vent window!

SUVIT filtergraph package (FGP)!
SUVIT spectro-polarimeter package (SPP)!

Bus module!
Middle gain antenna (MGA)

X-ray imaging telescope (XIT)!

+Y+Y

Solar UV-Vis-IR telescope (SUVIT) +X
+Z*+Z*+X* Boresight direction of the telescopes

Major concerns from the view point of system design- Total weight (The weight of the mission instruments is increasing...)- Pointing stability performance (Angular resolution of the telescopes are 2-5 times
higher than the that of similar telescopes onboard Hinode)- Cost (There is a very strict cost-cap for middle-size scientific satellite in JAXA)

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Technical Features of Solar-C Spacecraft #1SOLAR

Comparison between requirements andachievementsHigh Pointing Stability Requirement
Masada, Katsukawa et al.Requirements on the pointing stability of Solar-C
Instruments

SUVIT

EUVST

NormalincidenceXITXIT

Grazing
incidence

Time scale

Requirements(θx/θy)Unit
Orbital variation

1 sec

0.02

arcsec 3σ

10 sec0.02

arcsec 3σ

1 hour2

arcsec 0-p

Mission life

20

arcsec 0-p

0.5 sec

0.1

arcsec 3σ

5 sec

0.1

arcsec 3σ

1 hour2

arcsec 0-p

Mission life

32

arcsec 0-p

1 sec

0.1

arcsec 3σ

10 sec0.1

arcsec 3σ

1 sec

0.3

arcsec 3σ

1 min0.7arcsec 3σ1 hour8arcsec 0-p

Mission life

32

arcsec 0-p

Requirements(θz)

Instruments

Time scale

SUVIT

1 hour50

arcsec 0-p

EUVST

1 hour100

arcsec 0-p

Normalincidence

1 hour50

arcsec 0-p

Grazing
incidence

1 hour100

EIS/XRTrequirements
0.1 Hz hump

SOT
requirements

Comparison between the SOLAR-C
requirements and Hinode achievements
Masada, Katsukawa et al. &+,(7(.(/54,/,/0'(?!>!<= (3(,4$446.('4,/640,'$-
7,%3$5,0/450&0/7(35# 50# 1

+(611(3-(7(-40)5+(%$34,/',&$5(5+(3(26,3(.(/50)($&+,/4536.(/5$/'5+(-08(3-(7(-40)5+(%$34,/',&$5(5+($--08$%-(-(7(-  0)5+(3(26,3(.(/54)03($&+4063&(0)5+(7,%3$5,0/4


XIT (GI)

XIT (NI) / EUVST
SUVIT

Unit

arcsec 0-p&+,(7(.(/54,/,/0'(?!>!<= 
(3(,4$446.('4,/640,'$-
On-orbit
performance of the pointing
stability in Hinode
7,%3$5,0/450&0/7(35# 50# +(611(3-(7(-40)5+(%$34,/',&$5(5+(3(26,3(.(/50)($&+,/4536.(/5incomparison with the requirementsof Hinode (top)$/'5+(-08(3-(7(-40)5+(%$34,/',&$5(5+($--08$%-(-(7(- 0)5+(3(26,3(.(/54)03($&+4063&(0)5+(7,%3$5,0/4
and Solar-C (bottom).
1

Required pointing stability is 2 - 5 times
higher than the that of similar telescopes
onboard Hinode.

5Direction of system design for achieving
the required pointing stability
Science requirementFrequency band
Telescopes / Spacecraft system

SUVIT

EUVST

XIT

Absolute pointing
accuracy

SOLAR

Pointing stability performanceLong term

DC to 0.001 Hz

Mid. Term
0.001 to 0.01 Hz

Low frequency band

Short term
0.01 to 10 Hz (TBD)

over 10Hz

Mid. Frequency band

High frequency band

Image
Imagestabilization
stabilizationsystemsystemininSUVIT
SOT
Image stabilization system in EUVST

the range of band widthof each control system isunder consideration...

Image stabilization system in XIT

AOCSSpacecraft system

Low thermal distortionstructure
Orbit Selection

Passive isolator
for RWs
Reduction of the disturbance sources(Non-mechanical gyroscope, etc)

Ideas for the design improvements based on Hinode’s experiences:- to select an orbit with less orbital variation and low disturbance environment.
- to install passive isolator system for RWs.
- to adopt non-mechanical gyroscopes for IRU.
- to have dedicated image stabilization system inside each telescope.

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Workflow of the disturbance management
activities (concept)

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Apply the methodology established in the development of the Hinode spacecraft withfurther improvements. Detailed verification scenario on the pointing stability performanceis under consideration.....
Development phase

Activities in the disturbance managementPreliminary analysis

Conceptual design- Rigid body / Attitude and pointing control system model+ Upper limit of the transfer characteristics- Measurement of actual disturbance of each source(BBM or similar product)

Detailed analysisBasic design- Satellite finite element model (FEM) + Coupled analysis(structure/optical/attitude and pointing control system)- Measurement of actual disturbance of each source(BBM/EM or similar product)

Test to evaluate transfer
characteristics of disturbance 
Detailed design
Design tuning

Manufacturing

Vibration exciter+ Satellite MTM

Output of the disturbance management• System design concept• Preliminary error budget• Specification of the control system andeach component• Pointing error analysis• Requirement/allowable criterion foreach internal disturbance item
• Verification and ground test plan• Estimation of pointing stabilization
performance

Improvement of disturbance source
and/or transfer characteristics

Disturbance response test

Actual disturbance source+ Satellite PFM• Verification of pointing stabilization
performance

LaunchOperation
Operation

On-orbit wideband measurement withthe correlation tracker (CT)• Estimation of actual pointing error• On-orbit performance evaluation

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Technical Features of Solar-C Spacecraft #2

Operation in Geo-synchronous orbit

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Trade-off studies on the operation orbit have been performed for several orbit candidates.
The viewpoints considered in the trade-off study;- High-rate science telemetry downlink (100 Gbytes/day)- High pointing stability (less orbital variation, low disturbance environment)- Thermal design (especially the thermal design of the SUVIT telescope)- Continuity of observation (shorter eclipse seasons)- Real-time science operation- Feasibility of the launch by H2A (type 202) or next generation launch vehicle
Baseline Orbit : Inclined geo-synchronous orbit (GSO)‐$ Similar to Solar Dynamic Observatory (NASA)‐$ Altitude: 36,000 km, inclination: 28.5 deg (baseline), period: 1 day

23 days
34 days
47 daysEclipse duration in the case of GSO

Visibility of the spacecraft from USC, Japan8

Concerns from orbit trade-off study
Sun-synchronous polar orbit(SSO)

High-rate sciencetelemetry downlink

Need to use Ka-band (~80Mbps,QPSK), which capability is newly
required to ground station in polar
region. 15 passes per day.
Need further survey on the groundstation availability.

Inclined geosynchronous orbit(GSO)

Geo-stationary earth orbit(GEO)

X-band (16QAM) is recommended but required ~8hrs contact with USC
34m (24Mbps) or Katsuura (12Mbps max). Additional ESA and NASAstations are helpful to ensure the 8hr requirement.Ka-band (80Mbps,QPSK) requires only 3 hours contact, but concerns are rain attenuation
due to humid climate in Japan and new development of station facility.

SOLAR
Halo orbit at the Lagrangian point
L1X-band (24Mbps, 16QAM) is possible
with 30W transmitter and Ka-band
possible with the rate higher than
24Mbps, but it needs ~8hr occupation ofthe 64m Usuda antenna. Severe time
c o n fl i c t w i t h B e p i C o l u m b o M M O
operations will make this option difficult.
NASA DSN supports may be an option
for Ka-band.

High pointing stability

The Ka-band antenna is quickly The motion speed of the antenna is slow (~30 deg/hr), but need to An angular motion of the antenna is very
moved to direct to ground station evaluate the micro-vibration level.
slow (~15 deg/month), but need to(>100 deg/10min) every 98min,evaluate the micro-vibration level.
affecting the pointing stability.

Thermal design

The orbital variation by the periodic The orbital variation is very small and stable, providing a stable thermal environment.
infrared radiation from the earth is Earth albedo and infrared radiation is very small, making the thermal design easier, especially for the SUVIT mainconcerned on the high resolution telescope.
performance of telescopes.
Earth albedo and infrared radiation
enlarges the size of the radiators forcooling mirror and detectors.

Continuity of observation

An eclipse season (for 3 months) Eclipse seasons with about 70
with about 20 minutes (at maximum) minutes (at maximum) per day andduration every 98 minutes.
it continues 20 days, twice everyyear. Related potential demerit is toneed a large capacity of the battery.
Orbit maintenance maneuvers(twice a year) are needed everyyear.

Real-time science
operation

Real-time access is possible but it is Real-time access can be arranged in the duration of pass (8 hr maximum).
restricted to 10 minutes every 98
minutes for a polar ground station.
No critical issue.

Feasibility of the launch
by H2A 202 rocket

The spacecraft mass constraint is a
critical issue. In the preliminary
study, the total mass of thespacecraft is about 4 tons and it isalmost upper limitation of the H2A
202 rocket capability.

Eclipse seasons with about 70 No eclipse season.
minutes (at maximum) per day and it Orbit maintenance maneuvers (severalcontinues 50 days, twice every year. times a year) are needed every year.
Same potential demerit as SGO
exists.
Orbit maintenance maneuvers (twicea year) are needed every year.

More propellant (about 200kg)
compared with GSO is required.
Current mass estimation has
negative margin for H2A 202capability.

No critical issue on the mass limitation of
the spacecraft.
Detailed study on the orbit design
including transfer orbit will be needed.

9Technical Features of Solar-C Spacecraft #3

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High-speed mission data downlink

Requirement on the data rate of the mission data from the mission instruments (after compression)
Instrument

Estimated data rate (Mbps)
Standard

Burst

Spectro-polarimeter (SP)

1.2

16.0

Filtergraph (FG)

1.3

32.0

1.7

5.1

XIT-NI

1.2

48.0

XIT-GI

2.5

26.0

Total data rate (including XIT-PC)

7.9

127.1

Total data rate (no XIT-PC)

5.4

101.1

SUVIT
EUVST
XIT

Data volume produced each day (GByte)〜100

High speed dedicated telemetry channel is absolutely required for Solar-C Mission.
Baseline: X-band telemetry system‐ up to 24Mbps with 16QAM modulation‐ X-band/16QAM technology is used in several Japanese S/Cs
- 8 hours telemetry downlink is required.

Option: Ka-band telemetry system‐ up to 80Mbps with QPSK modulation‐ The usage of Ka-band downlink is new exploration in Japan. The installationof the Ka-band receiving systems at the ground stations is needed.
- Technical concern is the attenuation of signals due to bad weather in Japan.
- International collaboration on the ground station support should be beneficial.

10SummarySOLAR• Current baseline of the Solar-C spacecraft system design isdescribed. (Just 5 minutes overview! For detailed information,please refer to MPD chapter 5.)• The system requirements and the spacecraft design are NOT
solid in this phase (pre-phase A). There are still some flexibility
and options in the system design (e.g. the frequency of themission telemetry).
• Any feedback is definitely helpful for the definition of thesystem requirements.

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