8.

Threshold science mission (draft) [1-4-3]
最低限達成しなければならない科学 output のドラフト

The threshold science mission shall execute the observations necessary to: trace the mass andenergy flows through the atmosphere from the chromosphere into the corona; study energy releasein small-scale transients (e.g., explosive events, bright points); and determine the properties ofwaves throughout the atmosphere.
The threshold science mission will meet all of the Solar-C_EUVST requirements with the following
exceptions:• Spatial Resolution: An angular resolution better than 0.8”, which is two times worse than thetarget spatial resolution.
• Temperature Coverage: At least three wavelength bands to measure line intensities andderive electron densities and temperatures in the thermal coverage from the chromosphere to
the flaring corona.
• Field of View: Obtain repeated raster scans of representative targets in the quiet Sun, active
regions, and coronal holes on the solar disk over a limited field of view [100” (scan range) ×100” (along the slit)].
• Effective Area: Reduced throughput by factor 4 that still has more than twice larger effective
area than past and current instruments• Co-alignment: Shall obtain images by the slit jaw imaging camera in at least one wavelength
for co-alignment.
The minimum spatial resolution is still better than that achieved so far with past and current ongoing missions covering coronal temperatures; the spectroscopic data from Hinode/EIS has 2”-3”,
and from SoHO/SUMER has 1.6”, and the coronal images from SDO/AIA has 1.6” (Table 6.4).
Seamless temperature coverage is still maintained with three cameras, although some spectral lines
would not be available. Velocity diagnostic performance, i.e., Doppler shifts and non-thermal
broadening, is kept to trace mass and energy flows through the atmosphere. The target region could
be restricted to inside the solar disk due to pointing limitations, but some of specific tasks relating
to the solar wind in objective I-3 would be affected. The minimum field of view of a nominal rasterscan allows us to perform many of the observing modes defined in Table 6.7. Small-scale transients,
such as nanoflares, microflares, various types of jets, bright points, and explosive events, can be
captured within the limited field of view. High temporal resolution is required to detect the Alfvenic
signatures (1 sec exposure per 1” slit scan step to follow the Alfven speed at the corona, i.e., ~1,000km/s). If the effective area is reduced by factor of 4, the signatures can be followed by slit scanningwith exposure 1.6 sec and 1.6” step (equivalent to the spatial resolution of SUMER and AIA). Slit
jaw imaging at one wavelength will allow for co-alignment with ground based observations.
9.

Gross characterization of space environment [3-2-3]
ミッションを実施する宇宙環境

The proposed mission concept requires the achievement of sub-arcsec spatial resolution ofspectroscopic observations for spatially resolving elemental magnetic structures in the upper solar
atmosphere. It also requires continuous viewing of the Sun from the orbit for investigating the
temporal evolution of solar phenomena on timescales longer than at least one hour. Experience fromprevious spectroscopic telescopes on SoHO and on Hinode has shown that a stable temperatureenvironment is mandatory to reach optimum optical performance characteristics. There are threeorbit candidates for a stable temperature environment with continuous viewing of the Sun: 1) L1
point, where SoHO is operated, 2) sun synchronous orbit (SSO), into which Hinode, TRACE and
IRIS are installed, and 3) geosynchronous orbit (GSO), where SDO is currently operational. The
Epsilon rocket vehicle is used for the orbit installation of the proposed mission, which can install a
40/100590 kg spacecraft into SSO in 500-650 km, as described in section 4.2. It can bring a very small
mass to GSO and L1. Thus, SSO is the only option for the proposed mission. In SSO, the adverse
effects of Earthshine and Albedo on the instruments need to be minimized by optimal radiatorplacement and a specific orientation on the instrument. Moreover, compensation heaters will beneeded to dampen short-term temperature excursions, as determined from the Hinode experience.
The following three environments will be considered when the orbit height of the SSO isdetermined in the design phase. 1) The proposed launch timing is at the solar maximum phase, inwhich the atmospheric density is increased in the orbit, giving larger atmospheric drag to thespacecraft. A sufficient height of at least 550 km is required for a longer duration mission. 2) SolarUV light is slightly absorbed by the Earth’s atmosphere. It is better to reduce the period of theeclipse season in UV light. To achieve this, a height higher than 600 km is required (see Section
15.4.1). 3) Radiation effects (snow noise) appear in the images when the spacecraft travels throughthe South Atlantic Anomaly (SAA), and in high latitude Aurora zone during magnetic storms.
Radiation effects are more significant at higher altitude.
9.1.

Identification of planetary protection requirements [14-2-1]
Planetary protection の認識

Not applicable.
10.

Draft requirements on spacecraft system [4-2-1]
宇宙機（衛星・探査機）システムへの設計パラメータ・性能要求のドラフト

Performance requirements to the spacecraft system with design parameters are described here.
(a) Structure design• The proposed EUVST instrument shall be accommodated on the bus module, and the entire
envelope of the spacecraft shall be fitted within the Epsilon fairing envelope.
• The mounting for the EUVST instrument will be designed so that the optical performance of
the EUVST instrument should not be affected by the thermal deformation of the bus module.
• The ultra fine sun sensor (UFSS), for attitude control system, shall be mounted either on the
EUVST instrument or the mission tower for mounting the EUVST instrument.
• The spacecraft design shall satisfy the requirements on the interface with the Epsilon rocket,
including the mechanical interface at the rocket separation, payload environmental condition,
static balance and 1st natural frequency, which are specified in Epsilon Rocket Users’ Manual(NC, March 2016).
(b) Spacecraft orbit and propulsion• The spacecraft orbit (and its parameters, such as altitude, inclination and local time at thedescending node) during science operations shall be chosen to meet a) the requirement on thecontinuous viewing of the Sun (section 9), b) the requirement on the data transfer (section
6.4), and c) the requirement on stable temperature environment for achieving optimum optical
performance characteristics of the mission instrument (section 9). In any low Earth orbit, suchas sun synchronous orbit (SSO), the orbit parameters also need to be considered for the threeenvironment conditions discussed in section 9.
• The propulsion system, i.e., RCS (reaction control system), shall be onboard to perform theorbit maintenance, if the total duration of non-eclipse season (continuous viewing of the Sun)is reduced to less than 8 months per year within three years after the launch. The accuracy of
the initial orbit to the sun synchronous condition and the decrease of the altitude due to air
41/100

drag at the time of the solar maximum are considered. When the spacecraft is installed into
SSO by Epsilon rocket with PBS (Post Boost Stage), the accuracy (3s) in the altitude andinclination of the installed orbit is ±20 km and ±0.2 deg, respectively, according to Epsilon
Rocket Users’ Manual (NC, March 2016).
• The risks for producing debris shall be considered, including 25 years deorbit requirement.
• The eclipse season shall be taken into consideration for best coordinating with ground-based
observatories. Not having the eclipse season during the winter in the northern hemisphere is
the best option for coordinating with DKIST.
(c) Attitude determination and control• The spacecraft attitude shall be controlled by commands to point to any direction in theinertial reference system. This mode is mainly used in the initial phase without sun sensors.
• During the observations, the spacecraft attitude shall be stabilized in three axes with its Z-axis
pointed to the Sun.
• During the observations, the spacecraft Z-axis shall be directed to any position on the solar
disk as well as in the off-limb region (less than 9 arcmin or 0.15 deg) above the limb. The
positions are specified on a sun-oriented coordinate.
• The absolute accuracy of the pointing on the solar disk shall be defined as follows to easily
capture the target of interest within the EUVST medium-sized field of view:qx, qy: 20 arcsec (0-p)(the angle around two axes (X and Y-axes) perpendicular to the Z-axis)
qz: 100 arcsec (0-p)(the angle around the Z-axis).

•

•
•

•

The scientific operation team will determine a more accurate relationship between the attitudecommand and the actual pointing of the EUVST instrument by analyzing the scientific dataacquired with the EUVST on orbit.
As a baseline, the spacecraft shall track a region on the solar surface by correcting for solar
rotation during the observations. For each tracked target, the angular velocity around therotation axis of the Sun can be specified; The parameters for specifying the angular velocity
are prepared in the planning on the ground and uploaded to the spacecraft before theobservations. The onboard attitude control computer calculates the tracking motion dependingon the solar latitude and longitude and updates the spacecraft pointing in real time. The
maximum duration of the continuous tracking is two weeks.
The other tracking/observing mode is the spacecraft pointing to a fixed position on the solar
disk and above the limb. The fixed pointing is particularly important when observing theregion above the limb.
As a baseline, the Y-axis, which is parallel to the slit of the EUVST, is directed toward thesolar north. In addition to the baseline orientation, the spacecraft shall have a capability forroll control around the Z-axis, i.e., a capability to tilt the Y-axis from the solar north. The
purpose of this capability is to increase the observing cadence of spectroscopic observations,by aligning the slit direction to a specific target, e.g., the coronal loop of an active region ofinterest. It is noted that a reduced performance on the mission data telemetry speed is
acceptable in the roll orientation far from the baseline orientation.
The attitude determination is made onboard for attitude control. It uses the attitude knowledge
from attitude sensors, including Ultra Fine Sun Sensor (UFSS), Inertia Reference Unit (IRU),
and Star Tracker.

42/100• The calculation of the orbit phase, i.e., where the spacecraft flies, is needed onboard for the
EUVST instrumentation (note in Table 10.1). It is also needed for scientific data analysis.
Sufficient accuracy has been achieved according to Hinode experience.
• The pointing stability specified in Tables 10.1 and 10.2 and Figure 10.1 shall be achieved witha combination of the attitude control, the image motion compensation system in the EUVST
instrument (Section 15.1.4), and the structure. In general, the attitude control takes care ofjitter in the low frequency region, the structure avoids the jitter (micro-vibration) in the highfrequency region, and the image motion compensation system compensates the jitter inbetween the low and high frequency regions.
Time
5 sec
1 hr

Table 10.1: Pointing stability1): qx, qy
Requirements
Unit
Note
0.2
Based on the budget allocation for thearcsec (3s)spatial resolution (see Table 15.4).
2.0 2)
arcsec (0-p)
Drift smaller than the narrow FOV for the
fast scan shown in Sections 5.1 and 6.4.1.

1) Only for out of the eclipse season.
2) Pointing may slowly drifts with the orbital phase due to the thermal deformation. Building on the
IRIS's heritage on the correction by an orbital wobble table (De Pontieu et al. 2014), this orbital drift
will be suppressed to the goal level depicted by the red dashed line in Figure 10.1 (0.4 arcsec [0-p])
by a look-up table for correcting the tilt of the primary mirror.

Figure 10.1: Pointing stability (qx, qy) requirements in the frequency domain. See note (2) inTable 10.1 for the red dashed line. Assuming sinusoidal oscillations, Δθ 0 − 𝑝𝑝 = 2 3 Δ𝜃𝜃 3𝜎𝜎
was used to convert ∆θ 3𝜎𝜎 to ∆𝜃𝜃 0 − 𝑝𝑝 .
Time
1 hr

Table 10.2: Pointing stability1): qz
Requirements
Unit
Note
100
arcsec (0-p) ~ 1 pixel at the edge of FOV

1) Only for out of the eclipse season.

• When a pointing change is commanded, the pointing stability specified in the previous bullet
shall be achieved in the timescale less than 5 minutes after the pointing change.
• In the contingency of the attitude control, the spacecraft attitude mode shall immediately bechanged to safe-hold, in which the spacecraft attitude is kept in a safe orientation in respect tothe solar direction. The safe orientation shall be defined in the design phase. In the safeorientation, the mirror in the EUVST instrument focuses the solar light into a robust area in
the instrument, not giving any damage to the telescope structure.
• The spacecraft shall control its attitude properly to allow the EUVST instrument to performthe observations even during solar eclipses (the Moon passage between the Sun and Earth).
43/100(d) Thermal design• The thermal design shall guarantee the requirements on the temperature of onboard parts andcomponents both in operational and survival (non-operational) modes.
• The external surface of the structures in the bus module and the EUVST instrument is coveredby the MLI (multilayer insulation) to minimize the radiative coupling with the external
environment.
• The exception is the radiators necessary for the dissipation of the heat from onboard opticaland electrical components.
• In the SSO orbit, the thermal design shall carefully investigate how significantly the Earth’s
Albedo influences the thermal stability, especially the orbital variation in optical performanceof the EUVST instrument.
• The eclipse season is the thermally unstable period in the SSO orbit. Especially in the night
period and a certain time after the sunrise at each orbit, it is acceptable that the EUVST
performance, such as the spatial resolution, would be partially degraded.
(e) Power• The spacecraft will have solar array paddles (SAP). Choosing the SAP configuration shall be
determined from the power budget from the mission payloads. The upper limit of the power
supply capability during day periods is 1,000W for the entire spacecraft and payloads.
• The standard bus has a Li-ion battery (50Ah), which supplies the electric power to thespacecraft components in the period when the sufficient power is not generated by the solar
array paddles.
• During the launch phase, the battery shall supply the power to the spacecraft systemcomponents until the sun acquisition is completed.
• During the observation phase, the eclipse season happens about four months every year. In the
SSO orbit, the duration of the eclipse is 15-20 minutes in each orbit.
• During the observation phase, solar eclipses (the moon passage in front of the Sun) are
observed a few times every year. The battery may be used when insufficient power isgenerated by the solar array paddle.
• The battery Depth of Discharge (DOD) shall not exceed 25% in solar pointing and safe-holdattitudes, and 80% in case of transient battery mode, such as during launch and solar eclipses.
(f) Command and data handling• The science requirements on the mission data rate are specified in section 6.4.2.
• The spacecraft shall allow the EUVST instrumentation to perform continuous observations for
24 hours every day with the average telemetry rate of 0.72 Mbps.
• To download the science data acquired by EUVST, a high-speed data downlink system (Xband 8 Mbps as the baseline, or Ka-band) shall be onboard as a part of the mission payload.
With 15 downlink passes (duration: 10 min each) per day, 72 Gbits can be downloaded to theground in each day, making it possible to perform continuous observations with an averagerate of 0.8 Mbps for 24 hours.
• The spacecraft shall have a data recorder (2 GBytes) for recording house-keeping data without
losing the data. The house-keeping data will be downlinked via the standard S-band system(2Mbps maximum).
• In addition, a mass memory is required to store the science data from the EUVST instrument.
The minimum data storage for the science data is 20 Gbits (section 6.4.2), but a larger volume
is preferable for flexible observations. The proposed concept is that this memory is located in
the EUVST instrument.
44/100• The spacecraft shall handle the science data without losing the data, even when the data is
rapidly generated with a burst rate (the highest rate) specified in section 6.4.2.
• The time information shall be attached to the science data from the EUVST instrument. The
accuracy of the time information shall be better than 0.1 sec, after the ground post-processing.
• For downlinking the science data, the mission plans to use the ground stations not only inJapan but also those supported by other space agencies (similar to the case of Hinode). The
compatibility shall be confirmed during the ground testing phase.
(g) Mission duration• The science team requests two years as the normal operation phase after the launch, early
operation and verification phase, and consumables to support an additional three years ofextended operations.
(h) In-flight fault management• The spacecraft shall have the failure detection and recovery (FDIR) logic onboard.
• When the FDIR detects anomalies specified in the design, the spacecraft shall issue onboard
commands to turn off the components that are not needed to keep the spacecraft functional in
minimum and save the power.
• When the spacecraft detects anomalies critical to maintain the spacecraft pointing, thespacecraft shall change the attitude mode automatically to a safe-hold mode.
10.1. Comparisons to similar spacecraft systems[4-2-2][4-3-1]
過去の類似なシステムとの比較
Previous similar spacecraft systems are compared with the proposed spacecraft system and theunique requirements of the proposed system are highlighted.
a) Hinode (Solar-B), launched on September 2006

If the system functions and performance achieved in the Hinode spacecraft are available, they meetmost of the requirements in the proposed mission.
The satellite Hinode (Kosugi et al. 2007) was launched by the seventh M-V launch vehicle into an
elliptical polar orbit (perigee 280 km, apogee 686 km). It boosted its perigee and controlled theplane of the orbit with its own thrusters to acquire a circular, sun-synchronous, polar orbit of about
680 km attitude, 98.1 deg inclination, and 98 min period. With this orbit, Hinode has been
observing the Sun continuously for a duration of nine months every year. The height of the Hinodespacecraft is 4,000 mm and the weight about 900 kg, including 130 kg thruster gas, meaning thatthe proposed Solar-C_EUVST spacecraft is lighter than Hinode by a factor of 1.8. Three telescopes
onboard Hinode are aligned in the Z-axis of the spacecraft and supported by an optical bench tube,
which is mounted on the bus module. The spacecraft is stabilized in three axes with its Z-axis
pointed to the Sun. The Y-axis is always directed toward the solar north, while a roll capability of
the Y-axis is proposed in the Solar-C_EUVST spacecraft. As a baseline, the spacecraft tracks aregion on the solar surface by correcting for solar rotation. The achieved stability of the Z-axis is0.3 arcsec (3s) in 10 s and 1 arcsec in 1 min, helping the telescopes to perform high spatial
resolution and highly precise measurements of the solar atmosphere. One of the telescopes, theSolar Optical Telescope (Tsuneta et al. 2008), with 50 cm aperture, has an image stabilization
system consisting of tip-tilt mirror mechanism and correlation tracker (Shimizu et al. 2008),
achieving the extremely stable pointing of the telescope in order of 0.03 arcsec (3s).
45/100For data downloads, besides the S-band, Hinode has a X-band (4Mbps, QPSK) for mission data
downlink, with an onboard data recorder (1GB). During the period when the X-band was fully
functional, X-band downlinks were carried out at almost all the passes over the north pole (15passes at Svalbard). In the proposed mission, more telemetry and data recorder capacity arerequested to enhance the science output.
b) Hisaki (Sprint-A), launched on September 2013

The ISAS small bus module developed for Hisaki and subsequent Epsilon’s M-class missions isplanned to adapt to the spacecraft system of the proposed mission. Comparing to the Hisaki system,some new additions will be required to the small bus module system.
Hisaki has a single EUV spectroscopic telescope with a height of about 2,850 mm and is mounted
on the top panel of the bus module. The overall layout of the proposed mission is quite similar tothat of Hisaki. The weight of Hisaki is 350 kg, which is lighter by 150 kg than the mass budget ofthe proposed mission. A thruster engine is not included in the Hisaki system. The spacecraft isstabilized in three axes with its Z-axis pointed to planets. The pointing stability is requested to be
2.5 arcsec (0-p), when the pointing signal from the guide camera is utilized in the attitude control.
Note that the ultra-fine sun sensor (UFSS) is used in the proposed mission. Hisaki has only S-band,
which is used not only for housekeeping but also for science data downlinks. The data recorder is
2GB.
c) ASNARO-1, launched on November 2014

ASNARO-1 is an earth observing satellite developed by Japan Space Systems (JSS) and NEC, witha high-resolution optical imaging instrument that can deliver imagery at a ground resolution ofunder 0.5 m in the panchromatic band and under 2 m for multispectral images. ASNARO satellites
use the same bus that Hisaki does. The total spacecraft mass is 495 kg with 250 kg allocated to thebus, 200 kg reserved for the payload, and a nominal propellant load of 45 kg.

The spacecraft bus is outfitted with a number of Attitude and Orbit Determination and Control
Systems to ensure accurate navigation and precise pointing capability. The exact achievement of thepointing stability is not publicly available, but high resolution images available on internet as wellas the initial study carried by NEC for this proposed mission indicates a good pointing stability.

The acquired images are recorded in a 120 GB Flush memory. The X-Band system achieves a data
rate of 800 Mbps using a 16-QM modulation. The antenna pointing mechanism allows precise
pointing of the X-Band antenna to track a ground station.
d) IRIS (NASA’s SMEX), launched on June 2013

Interface Region Imaging Spectrograph (IRIS, De Pontieu et al. 2014) is a NASA solar observation
satellite, funded through the Small Explorer program. The spacecraft consists of a satellite bus and aspectrometer built by Lockheed Martin. The mission has been operated in a sun-synchronous polar
orbit with the perigee of 623 km, providing eclipse-free observations for up to eight months per
year. The satellite's instrument is a 19-cm UV telescope that feeds a slit-based dual-bandpass
imaging spectrograph. IRIS obtains spectra in 4 UV passbands in 130 – 280 nm. A spatial resolutionof better than 0.5 arcsec is required and spectroscopic observations with 0.4 arcsec spatial resolution
are realized.

The IRIS mass is 183 kg with 87 kg for the instrument and 96 kg for the spacecraft bus. The
spacecraft bus, designed and delivered by Lockheed Martin Civil Space (formerly LM Sensing andExploration Systems), is a rigid design whose frame is machined from a single piece of aluminumwith honeycomb aluminum forward and aft decks. IRIS is three-axis stabilized. The attitude control
system (ACS) is gyroless, using two star trackers, four reaction wheels, coarse and digital sun
sensors, and a magnetometer. The instrument guide telescope provides a high-resolution pointing
46/100signal to the ACS during normal science operations. Magnetic-torque rods are used to manage themomentum of the reaction wheels, transferring energy to the Earth’s magnetic field as needed.
There is no propulsion system onboard. The ACS can point the IRIS telescope to any location on
the solar disk or above the limb within 21 arcminutes from the disk center, and roll the spacecraft(and thus, the spectrograph slit) up to ±90 deg (at 0 deg the slit is oriented parallel to N–S on theSun). IRIS is equipped with two omnidirectional S-band antennas for commands uplinking anddownlinking of engineering data, and an X-band antenna for downloading of science data. The Xband provides downlink at 15 Mbps including the overhead of Low Density Parity Checking(LDPC) 7/8 encoding. The effective downlink rate is 13 Mbps (excluding overhead) during up to 15
passes per day at Svalbard, as well as some passes in Alaska and Wallops.
From a survey of the spacecraft discussed above, we have identified new requirements for the
proposed missions concept. The requirements for the spacecraft system are based on the ISAS small
satellite bus.
• High pointing stability: The stability is already achieved in the Hinode spacecraft, but it isnecessary to achieve the equivalent stability on a small satellite bus. Similar to Hinode, thepointing signal from the ultra-fine sun sensor (UFSS) is used in the attitude control. The
micro-vibration control, which is important in stability at the high frequency range, is requiredin the development, although the method has been developed during the Hinode development.
• High mission telemetry rate: The ISAS small satellite bus has 2Mbps S-band telemetry system
only, which is insufficient for downloading the mission data. A higher telemetry system (Xband or Ka-band) with a large mass storage is needed as a part of the mission payload.
11.

Mission operation profile [6-2-1][6-3-1]
ミッション運用方針案

11.1. Science operations plan

The EUVST instrument will be operated to maximize science output and achieve the mission’s
science objectives. As the guideline for daily operations, the following items are considered in orderto prepare for a variety of programs (observing sequences) that optimize the science return:• Tunable parameters available for the instrument, such as the selection of spectral lines to berecorded, exposure duration, spatial binning, scanning range and steps, and map cadence,• The free space in the onboard data recorder available for recording and the downlink scheduleof the recorded data,• The solar conditions, such as the appearance of active regions and their activity level, coronalholes and prominences, and• The coordination with ground-based and other space-based observations.
The system of daily operations follows the format of the Hinode operations that have been used andimproved during its 11 years of operation. The mission is developed and operated underinternational collaboration, here we have the following essential conditions for designing theoperation system (Table 11.1).

Figure 11.1 describes the baseline concept of nominal science planning, with the list of key
elements in Table 11.2. This baseline concept is almost same as what has been done in Hinode
operations. The only exception is that the daily meeting will not be necessary during normal
operations -- the chief observer handles target selection and program selection and coordinates
collaborations, similar to IRIS operations, because there is only one instrument onboard.

47/100Table 11.1: Essential basis for designing science operations
1. Japanese leadership in
ISAS/JAXA takes responsibility for spacecraft operations.
satellite operations.
Daily commanding and housekeeping status checking will be
conducted at SSOC in ISAS.
2. Balanced contributions to
To ensure that researchers and graduate students can spend
science operations from
sufficient time on data analysis to maximize science return
Japan, US and Europe.
through refereed publications and PhD and MSc theses.
The team shall take care that too much burden is not placed on
graduate students who work in daily operations.
3. Related researchers andThey will be involved in the roster of chief observer, who
graduate students are highly makes daily plan and checks the quality of latest downlinked
encouraged to participate in data.
the science portion of daily The science portion of daily operations is a good educational
operations.
opportunity for graduate students to learn in depth spacecraft
operations and onboard instruments, broadening theirknowledge and enhancing their ability to communicate with
scientific proposers from over the world. This is crucial to
improve their employability.
4. Remote operations fromThe science portion of daily operations is conducted at ISAS athome institutes (after stablethe early phase of science operation phase.
daily operation is established)
A system will be considered to allow researchers to makeplanning remotely. Note that the remote planning has been usedfor Hinode operations since 2009.
Community

Proposers

Steering committee

Science Working Group (SWG)

Observation
proposal

Science Schedule
Coordinators (SSCs)

EUVST Team
Core Plans for science objectives

Operation
meetings

Chief Observer

Chief Observer

Advice
Review and approval
Scheduling

Proposed Plans

Monthly Meeting

Long-term scheduling

Weekly Meeting

Stawman scheduling for next weekEvery Friday AM

Daily Meeting

Detailed science & pointing schedulingfor the next uploadMon Sat AM

Observation planning

Prepare observing sequences

After verification, submit the commands via ISACS system

Real-time operations

Uplink the commands
Check spacecraft status

Figure 11.1: Basic concept of nominal science planning
48/100Table 11.2: Key required elements in science planning
1. Strike a balance between coreobservations and proposed
observations

2. Regularly review operations andscience activities
3. Operate on-orbit solar
observatory accessible fromscientists over the world
4. Plan the optimum observing
timeline for the changing solar
conditions

5. Quickly act for unpredicted flarepossibilities.

6. Verify the contents of thecommands before uplinking to
the spacecraft.

The EUVST science team has core observations which will be
prepared prior to launch to achieve the minimum scienceobjectives.
The team will also perform proposed observations, which aredefined by receiving proposals from scientists in researchcommunities over the world. This is useful for scientists to
conduct the observations coordinated with ground-based andspace-based facilities.
Science Working Group (SWG) reviews operations and science
activities regularly and give guidelines and advice to theoperations team.
Science Schedule Coordinators (SSCs) receives and reviews
proposals for observations, and schedules the approved proposed
observations.
The stream of monthly, weekly, and daily meetings provides
opportunities to discuss observation plans in the operation team.
Note that, similar to IRIS operations, a daily meeting will not benecessary during normal operations -- the chief observer handles
the target and program selection and coordinates collaborations,with a guidance from weekly meeting.
Considering variable solar condition, the frequency of timeline
upload will be changed from every day in weekday period (at theearly phase of the operations) to three-days interval (weekend, andat the late phase).
The team will have the list of priority for science objectives.
When a priority is given to flare observations, the team utilizes
any information for “flare watch” and makes a decision on lastminute changes in the observing region.
User-friendly chief-observers’ tools (COT) will be developed tomake the timeline and exposure setup. The tools (in COT and the
ISACS_PLN system) have protective functions to prevent
incorrect parameters from being sent to the spacecraft.

11.2. Telecommunications, Tracking, and Navigation

The standard S-band system in the ISAS small bus is used for telecommunication for commanding,housekeeping and ranging operations. USC34m, 20m, and KTU4 antennas as well as JAXA GN
antennas will be used for this purpose.
The standard S-band 2Mbps system is insufficient for downloading the scientific data. Thus, a highspeed telemetry system is requested for downlinking the mission data. The X-band QPSK 8Mbps,
which was used in Hitomi, could be the baseline, and Ka-band system with much higher rate might
be considered if NASA support is available. The mission data will be received at USC34m, 20m,
and KTU4 antennas in Japan as well as at additional oversea antennas supported by ESA and
NASA.
There are no requirements for the spacecraft to know the precise location in orbit in real time fornavigation. As for Hinode, the orbital information may be uploaded on a regular basis to the attitudecontrol system to recognize orbital events, such as night and day. For science purposes, there are no
requirements for the onboard instruments to know the location of the orbit, but scientists may wantto know the location on orbit for calibrating the mission data.
11.3. Ground systems and facilities

Figure 11.2 shows the overall structure of ground systems and facilities, mainly focusing on the
mission data flow. All the mission data, received at some ground stations, are delivered to the JAXA
49/100Sagamihara campus via a dedicated line or the internet. The overseas station data server at
Sagamihara retrieves the telemetry files from ESA and NASA stations with a time delay (typicallyless than a couple of hours), as performed for Hinode telemetry. In ISAS, the data distributor andstorage servers are used for data distribution.
All the telemetry data is archived after time stamping on the SIRIUS database. For science
purposes, the telemetry data is reformatted as FITS files; Quick Look files are automaticallygenerated immediately after the telemetry delivery, whereas the final reformatted files are createdafter confirming all the telemetry data achieved in SIRIUS. The final reformatted files are open to
the researchers via the DARTS system.

For the promotion of the Solar-C_EUVST science, a Japanese science center will be formed at
ISEE, Nagoya University, where analysis tools, calibration, and the computer environment for data
analysis are provided to researchers. The reformatted data are copied to science centers at US andEurope for the American and European communities.

KTU4(Secondary)Commanding
Monitoring

CMD
/HK

JAXA GN
Mission data

Mission data

USC34m/20m(Primary)

GN / KTU4 Ops(Tsukuba)

SLE GW CMD/HK

Mission data/HK

Planning

Mission data

Satellite telemetry database (SIRIUS)
Mission data/HK

Mission data
Quasi real-time

Chief observers’ tools
ISACS-PLN

Reformatting

Science data achieve (DARTS)
rsync

Chief observer

NASA stations(Alaska etc.)

Oversea station data server
SSOC

Sagamihara Ope Center (SSOC)

Timeline/Commands

ESA stations(Svalbard,
Troll etc.)

Mission data

CMD/HK/RNG

The lower left portion in Figure 11.2 is the section for planning the observation timeline andcommands, referring to the latest mission data available for science planning. We utilize theoperation tools developed in ISAS, including GSTOS (Generic Spacecraft Test and Operations
Software) for planning spacecraft commands and verifying the commands, SIB2 (Spacecraft
Information Base version 2) for definition of the commands and telemetry, and ISACS-DOC for
automatic monitoring of the telemetry status. SIB2 and GSTOS will be used for developing andtesting of the spacecraft and instruments. For science planning, PI-prepared chief observer’s tools(COT), evolved from Hinode’s planning tools, will be used to schedule the timeline and commands
for exposure sequence of the EUVST instrumentation. A pointing scheduling tool, based on
Hinode’s pointing tool, will be also required for safely preparing the sequence of attitude control
commands for pointing changes.

SearchingengineScience Center
Analysis @Nagoya
tools

Figure 11.2: Ground systems and facilities.
50/100rsync

Reformatted data

Science Center
Science Center

Research Community

US
Europe11.4. Launch and early operations, contingency

At the launch, the sequence of operations, such as the deployment of solar array paddles and theacquisition of the Sun, is critical for the survival of the spacecraft. The selection of ground stations
and their required functions (commanding, status monitoring, ranging and real-time connection
from SSOC) shall be carefully considered when the launch sequence is designed.
Close approach of space debris can be considered as one of contingent events, something relatively
frequent for the satellites in SSO. If the thruster system is on board, we are requested to avoid
possible collisions, by following the JAXA guideline concerning space debris risks (QNX-160019).
12.

Key technologies [8-2-1]
主要技術要素Table 12.1 shows the list of key technologies for the proposed mission. The items which require
new development are highlighted in beige, and are described in section 12.2.
Table 12.1: Key technologies and development items.
Category
MissionRequirement
Seamlesstemperature
coverageTechnology
Multilayer coating (Mo/Si+B4C)

CCD E2V 42 (EUV)
Intensified APS (UV)
Spatialresolution

Wavefront error of the mirror
Optical performance

Hinode/SOT,
Hinode/EIS,
SO/EUI
Hinode/OTA

Heat dump
Focus mechanism (1) Primary
mirrorFocus mechanism (2) Grating
Image stabilization (1) Tip-Tilt
mechanism for the primary
mirrorImage stabilization (2) Guide
telescope

Grating (EUV)
Grating (UV)
Temporalresolution

Heritage
Hinode/EIS
SoHO/EIT, SO/EUISTEREO/EUVI
Proba2/SWAP
SPICE (for B4C)
SDO/AIA,
Hinode/EIS
STEREO/SECCHIWISPR/PSP
SO/EUI, SO/SPICE
Hinode/OTA

Enlarged mirror diameter
Multilayer coating (Mo/Si+B4C)

51/100heritage from
MELCO, Solar-C
R&D
Hinode/EIS,SoHO/SUMER
Hinode/SOT, IRIS,
Solar-C R&D
Duplicate of IRIS
flight GT. Similar
to SDO/AIA,
GOES/SUVI,STEREO/SECCHI,
TRACE
Hinode/EIS,
SO/SPICE
Hinode/EIS,
SO/SPICE, CLASP
Hinode/OTA
SoHO/EIT,
STEREO/EUVI

Development item
Combining Mo/Si coating
and B4C coating

Ensure small thermal
deformation of the mirror
Ensure small optics
misalignment and thermal
deformation
Heat dump from the primary
mirror
Application of flight
heritage parts to themechanism.
Application of SOT heritageand Solar-C R&D
development to thismechanism.

Tolerance/manufacturing
accuracy of high-density
grating
Tolerance/manufacturing
accuracy of grating
Ensure high reflectivity

Spectral
resolution andcoverage

Grating (EUV)
Grating (UV)
Slit assembly (narrow slit on
silicon wafer)

Slit imaging
forcoalignment

APS imaging sensor
filter wheel mechanism

Band selection filters
Spacecraft

Pointing for
observingtargets

Pointing
stability

Data rate
Mission life

Operation&groundbased

Data transfer

Solar rotation tracking
Absolute pointing accuracy

Proba2/SWAP
SO/EUISPICE (for B4C)
Hinode/EIS,
SO/SPICE
Hinode/EIS,
SO/SPICE, CLASPVERISSoHO/SUMER
Hinode/EISWISPR/PSP
SO/EUI, SO/SPICE
Direct copies offlight mechanisms
on IRIS, SDO/HMI,
SDO/AIA, or
Hinode/SOT
Based on filters on
SDO/AIA or IRIS
Hinode/AOCS
HinodeOptimizing the grating
parameters
Optimizing the grating
parameters
2.5 um slit

FDIR and safe-hold attitude

Hinode/AOCS +OTA

Roll maneuver around the Zaxis

IRIS

Pointing stability (lowfrequency)

Hinode/UFSS

Pointing stability (highfrequency)
Large mass memory
High speed downlink (X-band
8Mbps)
Contamination control in UV

HinodeDetermine the relation oftelescope pointing and UFSS
pointing based on
observations.
Protect the telescope
hardware damage fromoblique incidence of thesolar light.
Design X-band
configuration (location,
number of antennas)
Re-development of UFSSdue to change in company
developing UFSS
Micro-vibration control(heritage from LM)
HitomiUse Hitomi X-band system

Data transfer from
Svalbard/Troll to SSOC

SoHO,
Hinode/SOT, EIS,
STEREO, SDO,
SO, Solar-C study
etc
HinodeImplementation issue

12.1. Assessment of uncertainties in technologies [8-2-1]
技術的不確定性の大きな要素のリスト化
12.1.1. Active primary mirror

The primary mirror of the telescope contains key technologies for determining a large part of
optical performance, such as the spatial resolution and throughput. The mirror is made of an ultralow expansion glass material, such as Zerodur or ULE, which has advantages to ensure very low
thermal deformation and realize very low surface micro-roughness for suppressing unwanted
scattered light. The heritage from our development on 50 cm diffraction limited mirror system for
Hinode/SOT can be extensively utilized, but the control of thermal deformation on the mirror
52/100surface figure will be one of the new key technologies because the mirror will be operated at a
relatively high temperature (80℃). With such a high temperature, the mechanism for supporting themirror may deform the surface figure, for example by a stress due to the difference of CTE between
the mirror and mounting metal pads bonded on the mirror. The high-quality mirror surface figure
will be achieved by polishing and verified at the room temperature (~23℃), but the thermal
deformation caused by a temperature change from ~23℃ to ~80℃ needs to be controlled below therequired level. With the design of a kinematic mount with three pads bonded on the side of themirror, a major aberration due the thermal deformation is a trefoil coma, which is controlled during
the polishing of the mirror. The mirrors of Hinode/SOT were well designed to minimize the thermal
deformation in their operational temperature up to 47℃ (Suematsu et al. 2008), which were verified
with extensive thermal-optical tests. A micro-roughness of <5 Å rms in the spatial frequency range
from 10-3 to 5x10-2 nm-1 will be achieved based on the experience in fabricating the 20cm Zerodur
mirrors for SDO/AIA (Lemen et al. 2012).
Another technology required for the mirror assembly is a focusing mechanism for obtaining
sharp/on focus images and a tip-tilt mechanism for stabilizing the image during the slit scanning. A
launch lock mechanism will be also required to hold the mirror during the launch. Each one of thementioned mechanisms has been developed (section 16.2.1) and applied to various space missions.
Our baseline is to utilize the heritage technologies, but whether or not these technologies areapplicable without major design changes largely depends on the design of the assembly.
12.1.2. Multilayer coating

A robust Mo/Si multilayer coating with a B4C capping layer is applied to the mirror surface. Thiscoating uses well-known and characterized materials and provides excellent performance over abroad wavelength range in the EUV-UV. The selection of a single mirror coating rather than two
separate tailored coatings minimizes the risk of wavefront reflections induced by a split-mirror
telescope. Moreover, a uniform coating (uniform reflectivity and emissivity) minimizes thermal
stresses throughout the mirror.
A design example of the mirror coating isshown in Figure 12.1, which has highnormal-incidence reflectance in the EUV
wavelength ranges 170 ‒ 215 Å and above
460 Å. The coating consists of a graded
Mo/Si multilayer capped with a 10 nm thick
B4C layer. The B4C capping layer on thetelescope mirror provides a reflectanceabove 30% in the wavelength range of the
LW channel. Although the individual
coatings have been qualified at temperatures
up to 200°C, the temperature and EUV
reflectivity characteristics of the coating
selected for the telescope mirror will
undergo extensive testing early in the designphase to verify its performance for flight.

Figure 12.1: Performance of a broadband Mo/Si
multilayer mirror coating (reflectance peak around
18.5nm) with 10nm B4C capping layer.

Tests of the B4C capping layer conducted for SPICE on Solar Orbiter showed that the coating must
be protected from incoming solar wind protons. The EUVST primary mirror will be protected by aproton deflector, consisting of two parallel plates providing an electrical field strong enough todeflect incoming protons from their path towards the primary mirror. Such a design was
successfully used in the SUMER spectrograph on SoHO and the SPICE spectrograph on Solar
Orbiter (Fludra et al. 2013).
53/100We should also note that, as mentioned in section 16.2.3, the alternative coatings may be available,
which would reduce the mirror temperature.
12.1.3. Gratings

The gratings (both LW and SW) are ruled onto a toroidal surface like that used in many prior solar
EUV spectrographs such as SoHO/CDS or Hinode/EIS, as well as several NASA sounding rocket
payloads (EUNIS and VERIS). However, instead of the traditional constant line spacing, EUVST
uses rulings with variable line-spacing (VLS) of the form developed by Kita et al. (1983) andHarada et al. (1995) in order to minimize optical aberrations. Spectrographs using VLS ruling on
spherical substrates have been flown on several astronomical satellite missions such as EUVE
(Hettrick et al. 1985) and ORFEUS (Hurwitz & Bowyer 1991). A VLS grating was used in the
Hubble Space Telescope, Cosmic Origins Spectrograph. The EUVST design combines theseprevious concepts by using a toroidal substrate with variable line-space grooves (TVLS grating).
This design provides high-quality performance and unprecedented plate scale for the EUVST. A
holographic, 2400 groves/mm TVLS grating has been fabricated by Horiba/Jobin Yvon (HJY) for
the Solar Orbiter/SPICE spectrograph (Fludra et al. 2013, Caldwell et al. 2017). TVLS gratings
have been flown on the NASA SUMI (holographic, HJY, West et al. 2000) and RAISE(mechanically ruled, Bach, Laurent et al. 2016) sounding rockets. As shown in Table 15.1, the SW
grating of EUVST has a relatively high groove density of 4200 lines/mm. The Hinode/EIS uses
4200 grooves/mm toroidal grating with constant line-spacing (TCLS) grating (fabricated by Zeiss).
The FUSE experiment used a 5700 lines/mm TVLS grating (fabricated by HJY). There is nofundamental difference between a TCLS and a TVLS, the substrate figure having no influence on
the holographic recording process. Nevertheless, we need to investigate achievable tolerance andmanufacturing accuracy by optimizing the grating parameters with vendors.
12.2. Assessment of heritages [10-2-1]
既存技術の活用条件
12.2.1. Detectors

The detector systems for the EUVST have extensive heritage as described in Table 12.1.

CCD focal plane assembly ‒ The EUV focal plane assembly will consist of two E2V 42-40
detectors. These detectors are butted together for minimal gap. The combined format is 2048x4096.
Back-thinning and coating of the CCD will achieve optimal EUV quantum efficiency and stability.
These detectors have been used extensively in the past. Current solar missions with E2V CCDs
include: Hinode/EIS, Hinode/XRT, Hinode/SOT, STEREO/Secchi and SDO. The analog drive
electronics will be nearly identical to the electronics used on previous missions and the expected
readout rate of 2Mpixels/s is well within the device capability. The camera front end electronics
development will incorporate an engineering model to reduce programmatic risk. The focal plane
incorporates a 1000 Å thick aluminum foil filter to reject the visible radiation. The camera assemblyis equipped with a focal plane shutter. The CCDs will be thermally isolated and strapped to an
external radiator to be cooled to below -40°C. Previous work for the Hinode/EIS instrument has
shown that this temperature will be sufficient to suppress dark current and warm pixels for the
expected on-orbit dose.

IAPS focal plane assembly ‒ The LW detectors consist of three independently operated intensified
active pixel sensor (APS) detectors. The focal plane assembly is similar to previous intensified focal
plane assemblies flown on SoHO/CDS, SO/SPICE, SO/Metis and SO/EUI. APS detectors are
utilized extensively on the Solar Orbiter mission in the remote sensing instrumentation. The APS
detector for EUVST is identical to the devices utilized on the SO/SoloHI. Residual devices from the
SoloHI lot run will be used for the EUVST. Each focal plane assembly will incorporate two
1920x2048 detectors with 20 µm equivalent pixels (at the micro-channel plate surface). The APS
54/100detectors are radiation hard (>100krad) and have excellent photometric performance (high gain:
19,000 electrons full well with read noise of ~3 electrons; low gain: 87,000 electrons full well witha read noise of ~15 electrons). The APS front end drive electronics will be nearly identical to thatbuilt for SO/SoloHI. The APS detectors will be coupled to the back end of the microchannel plate(MCP) intensifier by a 2:1 taper. The MCP intensifier provides high efficiency and blindness tosolar visible radiation such that no focal plane filters are needed. A filter wheel with an Al thin film,
order sorting filter will be used to observe the Ne VII 46.5nm emission line. The intensifiers arerobust and no difficulties with radiation are expected (O. Siegmund, private communication). The
intensifiers have relatively large area (8cmx4cm) but are well within the previous experience. The
MCP front faces will carry the selective photocathode coatings with CsI and/or KBr. The APS
detectors will be cooled to <-45°C. This requirement will be revisited during the initial designphase. To protect the photocathode and MCP, each detector assembly is equipped with a vacuum
door. The MCP and door assemblies have extensive flight heritage on a number of space missions.
Prior to delivery, the MCPs are scrubbed to minimize gain changes with accumulated use. After
scrubbing, the MCP housing is maintained at vacuum levels until launch using an ion pump. All
three IAPS assemblies share the same vacuum pumping line. To develop these detectors, a full
qualification model focal plane assembly and IAPS front end electronics will be incorporated into
the EUVST development plan.
12.2.2. Mechanisms

Table 12.2 shows the list of mechanisms which will be fabricated based on the heritages for the
spectrograph unit.
Table 12.2: Mechanisms in the spectrograph unit from heritage missions

Mechanism
Slit selection

Requirements and characteristics
5 slits on a linear translation or rotating mechanism

Grating focus

Linear grating translation for ~2 mm range with 10µm accuracy/repeatability.
CCD shutter
Direct driven shutter mechanism with opticalencoding. > 4x107 cycles.
IAPS detector door Re-closable door. Motor mechanism to insert Al
and order sorting filter to observe Ne VII 46.5nm line in second order.
filter mechanism
Filter wheel
Select a broadband filter in Slit-jaw imaging system.

Heritage
VERIS, SoHO/SUMERSoHO/UVCS, Hinode/EIS
SoHO/SUMER, Hinode/EIS
Hinode/EIS
OthersHST and other missions
SO/SPICE
TRACE, SDO/AIA

12.2.3. Guide Telescope

The Guide Telescope (GT) can be a near-identical duplicate of the IRIS GT that has worked
flawlessly on-orbit for 4 years. Similar GT’s have flown successfully on SDO/AIA, GOES/SUVI,STEREO/SECCHI and TRACE.
12.2.4. Pointing stability ~ UFSS

Ultra Fine Sun Sensor (UFSS) is a compact but high accuracy sensor for measuring the direction of
the Sun with the accuracy of 0.3” (s) for ±0.5 deg field of view. The component consists of two
orthogonally oriented sensors, where each one uses an optical reticle and a 1-dimensional CCD. An
onboard smart algorithm determines the solar direction with a high accuracy. The UFSS has been
perfectly working on the Hinode satellite. However, since this UFSS manufacturer no longer
manufactures the UFSS, ISAS is working with a new manufacturer to develop the second
generation of UFSS based on the design and knowledge acquired during Hinode (Section 16.2.4)
with support from the chief engineer who designed the original UFSS.
55/10012.2.5. Pointing stability ~ micro-vibration

Since high spatial resolution is at the heart of the proposed mission, achievement of the required
image stability given in Table 10.1 is of crucial importance. Although the attitude control systemand the image stabilizing system of EUVST contribute respectively to the image stability at low(<0.1Hz) and middle (0.1–10Hz) frequency ranges, image jitter at high frequency range (>10Hz) isbeyond their capabilities. Such high frequency image jitter is caused by the micro-vibration of
optical components in the telescope and the spectrometer. Therefore, the suppression of theiramplitudes below the allowable level is an issue. There are several factors that determine theamplitude of high frequency image motions; i.e., the level of disturbances of moving components inthe spacecraft (disturbance sources), influence of the micro-vibration of optical components on theimage motion (sensitivity) and transfer of mechanical disturbance from the disturbance sources to
optical components (transfer function).
The potential risk on the micro-vibration issue is in the fact that the accurate prediction of thetransfer function of the mechanical disturbance is difficult to determine during the design phase
especially in high frequency range in f >100Hz. For this reason, the amplitude of the microvibration of optical components is highly sensitive to the resonance frequency and the amplification
factor of the structure, and a slight difference between the mechanical model and the real hardware,
results in a significant difference in the final amplitude of the micro-vibration of optical components
between the prediction and the experiment. Therefore, in order to ensure the image stability to besatisfied in orbit, we need to combine the evaluation/control of disturbance with the Finite Element
Method (FEM) analysis in design phase and experiments using the real hardware in development
phase. A strategy plan for avoiding a significant amplification of micro-vibration is given in section
16.3.3.
The scheme and methodologies for controlling the micro-vibration were established in thedevelopment of Hinode, especially for the Solar Optical Telescope (SOT). After extensive works
including the evaluation of disturbance sources, prediction of the mechanical transfer functions by
FEM and their evaluation by measurements using MTM, and final end-to-end measurements,
Hinode/SOT finally achieved a superior image stability of 0.03” (3s) in orbit. We utilize thissuccessful heritage of Hinode at most in the development of EUVST. The overall scheme developedin Hinode can directly be applied to EUVST. Note that the required level of the image stability in
EUVST (0.2 arcsec (3s), Table 10.1) is about twice less tight compared to that of the Hinode/SOT(0.09 arcsec (3s)). However, there are some differences from the Hinode/SOT, i.e., the structure andthe performance of ACS of the standard bus module for the small satellite is significantly different
from those of Hinode, the image stabilizing system of EUVST will not be as fast as that of the
Hinode/SOT because of the large mass of the active mirror, and EUVST does not have a fast
imaging sensor in its focal plane that is available for the final end-to-end testing of the microvibration. These issues and revised methodologies, if necessary, will be investigated during thedesign phase.
12.2.6. Contamination control in UV

Contamination degrades the optical performance of the science payload by obscuration and scatter
due to the presence of particulates on the optical surfaces and by absorption of incoming light due
to materials deposited on optical surface originating from organic material outgassing from thepayload structures. Regarding the particulate contamination, the surface cleanliness of Level 300,
which is adopted in the Solar-B mission, is sufficient. The wavelength ranges that EUVST observes,the extreme-ultraviolet (EUV) and vacuum-ultraviolet (VUV) wavelength ranges, are the mostcontamination sensitive regions for the molecular contamination that degrades the key performanceof the high instrument throughput.
56/100

The requirement of the limiting molecular contamination needs to be defined for the longest
EUVST wavelengths near the hydrogen Lyman alpha line. The tolerable amount of thecontaminants on the optical surface is 250 ng cm-2 (so-called A/4 level [TBD] ) during any time of
the whole mission life. The budget is more severe than that in the Solar-B (Hinode) mission by anorder of magnitude. However, the EUVST international team members have experience indeveloping the science payloads on SoHO, IRIS, CLASP, and Solar Orbiter that work in the same
wavelength range, and the same methodology that has so far been successful will be applied to the
EUVST development program. The most critical issue in the EUVST is the molecular
contamination on the primary mirror for the photochemical deposition process under the direct solar
UV illumination. The contamination budget control, modelling, and verification activities that were
employed in the Solar-B program will be applied to the EUVST program as the baseline, with theefficient bake-out methodology adopted for UV instrumentation in the SoHO and Solar Orbiter
programs. A severe material selection and screening process will be employed that guarantees only
materials with proven space heritage will be used and the bake-out will be applied to all hardware.
12.3. Technical risk identification [7-2-1] and major risks [7-2-2]
技術リスク分析と主要リスク要因
Based on our knowledge at the current phase, we listed the potential risks that may hamper theachievement of the science objectives, in particular, in the performances of investigations, indevelopments of the instruments and flight systems, and in the mission operation profile and theground system (Table 12.3, in the next page). Risk assessment has been made based on the criteria
defined in Table 12.4 – 12.6. Five major-risk (“high” risk) level items (highlighted in Table 12.3)
have been identified at the current phase. They are related to the active primary mirror (2 items),
grating (1 item), which are key components in EUVST, and micro-vibration control (2 items). For
these items, risk level will be reduced to lower levels, by performing the feasibility studies incoming years, and micro-vibration evaluation at early phase of the development (Pre-Phase A to
Phase A).
Table 12.4: Definition of risk “level”

Impact of occurrence [High]Impact of occurrence [Med]Impact of occurrence [Low]
Level
Low
Med
High

Level
Low
Med
High

Possibility ofoccurrence[High]
High
High
Med

Possibility ofoccurrence
[Med]
High
Med
Low

Possibility ofoccurrence
[Low]
Med
Low
Low

Table 12.5: Definition of “Probability of occurrence” (%)

Definition
The possibility of the occurrence is low, and this risk will be typically avoided.
The possibility is not low, the action to avoid/mitigate the occurrence is needed.
The possibility is high, and there may be no substitute for avoiding the occurrence.

Table 12.6: Definition of “Impact” by the occurrence(Technology, Schedule, Cost are judged with their “or” relation)

Technology
Minor or no influence
Possible to accept ornext best available
Impossible to accept

Schedule
Major milestone is delayedin less than 1 month, or no
impact to major milestone.
Major milestone is delayed
1 month – 3 months.
Major milestone is delayed
longer than 3 months

57/100

Cost
Cost increase less than
10M yen or no increase.
Cost increase between
10M and 200M yen.
Cost increase larger than
200M yenTable 12.3: Risk evaluations


001

  
Performance

Performance002                  The grating performance cannot be
The tolerance in some grating parameters is too tightachieved, resulting in degradation ofand exceeds the manufacturer's capability of delivering High Med High High Low Med
spatial or wavelength resolution.
according to schedule.
Wavefront performance of the primary A thermal deformation is induced to the mirror surface
mirror after mounting does not meet the when the temperature is changed from the room
High Med High High Low Med
requirements in the allowable budget
temperature to operational temperature.
table.

The reflectivity and IR emissivity of multilayer coatingHigh
surface deviates from the design parameters.

Performance003

Performance,system004 development

005

Performance,
instrument
development

Med

High High

Low

Med

Low

Med

High

Low

Med

Med

Med

Med

Low

Low

The magnitude of microvibration caused by some MWs
Highpurchased exceeds the specification.

Low

Med

Med

Low

Low

Transfer function from the MWs to the optics in the
EUVST is larger than the assumption.

Med

High High

Low

Med

Low

Med

Med

Low

Low

Low

Med

Med

Low

Low

Low

Med

High

Low

Med

Low

Med

High

Low

Med

Low

Med

High

Low

Med

Low

Med

Med

Low

Low

Med

Med

Med

Low

Low

Low

Low

Med

Low

Low

Med

Med

Med

Low

Low

Low

Med

High

Low

Low

Low

Med

Med

Low

Low

Low

Low

Low

Low

Low

Low

Med

High

Low

Low

Med

High Med

Low

Low

Low

Low

Low

Low

Low

Low

Low

Med

Low

Low

Low

Low

Med

Low

Low

Med

Low

Low

Med

Low

Mis-alignment between the primary mirro The mis-alignment is caused by the launch environmentHighand spectrograph unit becomes large,
resulting in degradation in spatial
The thermal deformation is induced by the temperature
resolution and shift in observablechange on orbit.
Medwavelength bands.
Microvibration from MW degrades the
EUVST spatial resolution performance.

High

The image stabilization system does not A tilt angle sensor of the active primary mirror showsHigh
work properly, resulting in degradation of malfunction.
spatial resolution.
An actuator of the active primary mirror does not work
High
Guide telescope does not measure the solar direction
correctly (due to brighter sunlight, stray light etc).

High

Performance,
006 Operation

A focus mechanism does not work,
resulting in degradation of spatial
resolution.

The focus mechanism of the active primary mirror doesHigh
not work.
The focus mechanism of the grating does not work.
High

Performance,
007
Operation
Performance

Slit cannot be changed.

The mechanism for changing slit does not work properly.
High

008

Performance,
Operation
Instrument &system010
development009

Performance,
Operation
Instrument
012 development
011

Systemdevelopment
System014
development
Operation
015
013

016

Operation

CCD/ICCD performance is degraded.

The molecular contaminants are condensed on thecooled CCD surface, reducing the throughput
Med
performance.
The CCD cannot be cooled down to requiredMed
temperature.
The radiation damage increases the number of hot
Medpixels, giving noise in spectral profiles.
ICCD cannot make observations.
The door mechanism in the ICCD housing does not
High
open.
The throughput of the entire EUVST is The optical surfaces are contaminated by molecularMeddegraded, resulting in increased exposure materials.
duration and slower cadence.
The reflectivity of multilayer coating surface isMeddecreased.
No throughput, leading to the mission
The telescope door does not open.
Highfailure.
The stiffness of the EUVST system
If the damper is needed at the Bus I/F point for reducin
including the Bus I/F point is too low.
the transfer function, the 1st eigenmode may beHighsignificantly decreased.
The total mass exceeds the rocket launch The mass of EUVST is much larger than the allocation.
Medcapability.
UFSS does not meet the random/bias
The light level is largely deviated from the setpoint dueMednoise requirements
to improper calibration.
Numer of ground station passes insufficienDue to lack of international supports and budgetMedfor science
limitation.
A limited part of telemetry is occassionallyDue to last-minutes cancellation by station's reason, andmissed at ground stations.
due to low level of signal received at stations at low
Lowelevations.

12.4. Margins in major-risk items [8-2-2][8-3-1]
主要リスク要因に対するマージン
In Table 12.3, five items have been identified as “high” (major) risk items at the current phase. The
numbers below are identical to those used in Table 12.3. For these risk items, there are uncertainties
that will be addressed through extensive technical studies. By discussing the origins of each risk,
the approach for reducing risk levels and setting reasonable margins is achieved.
58/100• No. 001: The grating performance cannot be achieved, resulting in degradation of spatial or
wavelength resolution. The aim of the flight gratings is to achieve the requirements for the
spatial and spectral resolutions. The spot diagrams to be shown in Section 15.1.2 for the baseline
design are derived by ray tracing analysis with the grating parameters, and readily meet EUVST
imaging requirements. The imaging error budget (Table 15.4) includes an experience based
fabrication tolerance and generous margins. To reduce risk, the EUVST team will procure andtest prototype gratings prior to finalizing the flight grating procurement as described in Section
16.3.
• No. 002: Wavefront performance of the primary mirror after mounting does not meet therequirements in the allowable budget table for achieving the required spatial resolution. Two
factors will be considered. The first is minimizing the thermal deformation at a high operationaltemperature (+80°C). We designed Hinode’s 50 cm mirror assembly with thermal deformation of
15nm rms (~1/42l) at temperatures up to +47°C. A first thermal-mechanical analysis for theassembly will be carried in pre-phase A in order to know the magnitude of the deformation andset a large margin for the deformation. The second factor is the uncertainty in reflectivity of themirror coating in visible and IR. Small changes in the reflectivity cause changes in mirror
temperature, producing thermal deformation. This will be considered in the thermal analysis.
Additionally, we will study alternative mirror coating with higher reflectivity in the visible. The
thermal properties will be characterized with coating samples during phase A.
• No. 004: Micro-vibration from momentum wheels degrades the EUVST spatial resolution
performance. As described in section 16.2, a first micro-vibration analysis with the ISAS small
bus indicated that the momentum wheels have potential to give vibration to the EUVST at an
unacceptable level. In this analysis, a model fidelity is not high because of many assumptions,such as the structural model of the instrument and a large Quality factor as an uncertainty in the
structural properties. Efforts are required to reduce uncertainties in the evaluation.
• No. 012: The stiffness of the EUVST structure including the Bus interface point becomes low, if adamper is needed at the Bus I/F point to isolate the mission payload from momentum wheels’micro-vibration. The experimental study mentioned in No. 004 will determine the necessity ofthe damper. Depending on the mechanical design, the 1st eigenmode may be significantly
decreased to below the eigenmode frequency required by the rocket. A sufficient margin for the
eigenmode frequency should be checked in phase A study.
13.

Cost assessments from similar missions [21-2-1]
類似ミッションからコストの規模推定

The proposed Solar-C_EUVST mission consists of an UV spectroscopic telescope with the ISAS
small standard bus. Hisaki (SPRINT-A, see section 10.2), launched as the first satellite for series of
Epsilon rockets, has the similar configuration and thus the cost used for Hisaki can be used as thebasis for the proposed mission. The EUVST telescope is bigger than the telescope EXCEED on
Hisaki (28cm vs. 20cm diameter of the primary mirror) and requires higher spatial resolution, hence
increased cost.
Because of the cost cap, Japan will take responsible for a limited portion of the EUVST, with
substantial contributions from US and European countries. As described in section 15.10, the mainpart of the Japanese contribution is the telescope. Thus, as an initial guess for the telescope, we canrefer to the cost in the development of the Hinode Solar Optical Telescope (SOT; 50cm in diameter,
0.2 arcsec spatial resolution). It spent 4,000M yen in development of the entire optical telescope(so-called OTA), including the precise combination of 50cm primary mirror and secondary mirror,
collimator lens unit, and tip-tilt mirror mechanism and its control system (so-called CTM). The
59/100EUVST has the primary mirror only and its diameter is about half of the OTA primary mirror,
whereas the primary mirror becomes more complicated by including the focusing and tip-tilt
functions. Considering the differences, the initial target for the development of the EUVST
telescope would be around 3,000M yen.
The estimated total cost for the Hinode/EIS instrument was $30-40M and was dominated by thecost of the spectrograph. The projected cost for the EUVST spectrograph is approximately $60M.
The additional cost is driven by the increased complexity of the electronics, the additional cameras
and grating for the long wavelength channel, and the slit-jaw imaging system.
14.

Trade studies among mission concepts
異なるミッションコンセプトとの trade off
14.1. Trade space [12-2-1]
Trade-off study 対象

The instrument in the proposed mission concept is compared with other possibilities of mission
concept in this section. The following six steps have been taken to compare it with various
investigations and instrumentations, finally drawing the baseline concept.
(1) Direction of future research in the solar physics domain

There are many science objectives in the solar physics area which should be tackled in 2020s. It isbeneficial to compare the different investigations. This review has been carried out by the NGSPMSOT team.
(2) Comparisons with various types of instruments

The previous trade study will conclude one direction of investigations (observations of elementalscale processes). As briefly described in section 14.2, the Next Generation Solar Physics Mission’s
Science Objectives Team (NGSPM-SOT) report (a supplement to the proposal; Supplement A)
recommends the following instruments as the high priority instruments to be realized in mid 2020s
to the study of fundamental physical processes occurring in the solar atmosphere.
1.
2.
3.

0.3” coronal/TR spectrograph (T-09) (= EUVST)
0.2”-0.6” coronal imager (T-07)
0.1” chromospheric/photospheric magnetograph/spectrograph (T-01/04/05).

In addition, a photospheric/chromospheric magnetograph with a reduced size of telescope is also
considered for the trade-off study.
4.
0.2”-0.5” photospheric/chromospheric magnetograph (T-02)
Of these instruments, T-09 and T-01/04/05 are basically the same as the instruments that have been
considered in the Solar-C for JAXA’s strategic large mission opportunity. The EUVST in the
proposed mission concept is quite similar to T-09. These instruments are also compared with theSolar-C for JAXA’s strategic large mission.
(3) Instruments for spectroscopy

In section 6.2, UV spectroscopy (selected for the proposed mission concept) was compared withthree different spectroscopic methods (in soft X-rays and in visible light) from the viewpoint ofscientific performance. They are compared with each other as a trade-off study from programmatic
aspects, in addition to scientific trade-off.
60/100(4) Spacecraft and fairing accommodations for UV spectroscopy

The EUVST instrument was originally studied as a European instrument called LEMUR, which was
proposed to ESA as a Mission of Opportunity during the ESA 2010 call for an M-size mission(Teriaca et al. 2012). This proposal submission was evolved to the EUVST instrument in the SolarC proposal for JAXA’s strategic large mission in 2015. The length of this EUVST instrument wasvery long (4.3 m) to achieve a high spatial resolution. Thus, instrument designs of various lengths
are considered in the spacecraft layout to accommodate each of these designs in the given Epsilon
nose fairing volume. This trade off study determines the maximum length of the instrument.
(5) Survey of optical parameters

In order to define the baseline architecture of the EUVST instrument, we have made a parameter
survey in optical parameters, i.e., a trade study in optical parameters under the requirements fromthe science (section 5 and 6).
(6) Orbit and telemetry trade study

As described in section 9, three orbit candidates (Sun synchronous orbit, Geosynchronous orbit, andL1 point) exist for achieving sub-arcsec spatial resolution with continuous viewing of the Sun. The
high telemetry requirement, given in section 6.4, should be also considered in the trade-off study.
14.2. Trade studies [12-3-1, 5-3-2, and other various points]
トレードオフスタディ(1) Direction of future research in the solar physics domain

As briefly described in section 3.1, the NGSPM-SOT team has identified key scientific objectives inall the research areas related to solar physics, based on extensive reviews of the broad interests ofthe community with a public call for white papers of science objectives. The list is categorized as 3
scientific objectives, consisting of 17 science sub-objectives with 56 tasks (Table 3-1 in NGSPM’s
SOT report attached as Supplement A). Based on eight criteria (1. Relevance to NASA/JAXA/ESA
objectives, 2. Scientific impact on solar physics, 3. Scientific impact on other disciplines andresearch fields, 4. Inability of current/planned missions and ground-based facilities to accomplish
the objective, 5. Need for space observations, 6. Maturity of technology -- measurements canactually be made, 7. Maturity of methodology -- data can be inverted to get findings, 8. Widespreadinterest within the solar physics community), the team made evaluations and determined prioritiesof objectives and tasks for next generation solar physics missions to be realized in mid 2020s.
Although the full range of science objectives are compelling, supported by the community, andtechnically feasible within the next decade, no single mission could achieve all the identified
objectives. Thus, the team considered conceptual groupings of the objectives to clarify how themost science could be achieved within expected constraints and resources. They are two
avenues: physical mechanisms on elemental (small) scales, versus global processes affecting/involving large fractions of the solar interior and/or atmosphere and propagating into the solar
wind. This division is imperfect and subjective, as the processes are certainly not decoupled.
However, for the purposes of planning a mission to carry a suite of instruments, it is necessary toconsider the physical scales of the observables, and this ‘bimodal’ division is useful. It is the senseof the team that the better near-term opportunity for breakthrough discoveries by the NGSPM will
be achieved through the former avenue, probing beyond the state of our present spatial and temporal
resolution. Global, multi-vantage observations are scientifically compelling, but also more
technically challenging.

61/100Global, multi-vantage observations
For understanding global processes affecting/involving large fractions of the solar interior and/or
atmosphere and propagating into the solar wind, the team has identified the large benefit of vantage
points away from the Earth-Sun line, including both in or near-ecliptic stereo views or out-of-theecliptic high-latitude views. The science motivating these observations is compelling. Two major
gaps exist in our current observational capabilities that could be addressed by such observations.
First, although the STEREO and Ulysses missions moved away from Earth-Sun line, neither
observed the solar magnetic fields and flows needed for progress on several key science topics
enabled by these vantages. Solar Orbiter will, for the first time, obtain remote sensing observations
out of the ecliptic. However, like STEREO, Ulysses, and the Parker Solar Probe, it will not maintaina sustained vantage point -- a second requirement on several science goals. Sustained polar/highlatitude views (which, depending on the science goal, may range from multiple months to the solar
cycle) can be achieved by a constellation of spacecraft or by a single spacecraft on a carefully
selected orbit. Transformative progress would certainly be made possible by sustained, multivantage views. However, such an implementation concept exceeds the capability (cost cap, weightof the mission) of this Epsilon opportunity. Thus, this concept is out of our scope.

Observations of elemental-scale processes
Solar activity on the smallest length and time scales illuminates fundamental physical processes,
often with less complex interactions and configurations than larger-scale analogues. Game-changing
discoveries may be found on these elemental scales, such as the nature of magnetic stresses
resulting from braiding of magnetic field lines; the magnetic topology at the footpoints of spicules,
jets, and prominences; the transport of energy via Alfvén waves in flares; and the development andeffects of turbulence in all levels of the solar atmosphere. Most of the tasks for these science topics
make use of wavelengths that are not accessible from the ground or require long-duration
observations in a stable, seeing-free environment, necessitating observations from a space platform.
There is every reason to believe that a higher-resolution focus for NGSPM will be richscientifically. Given constraints on the capability (cost cap, weight of the mission), continuing toincrease resolution is both feasible and well-suited to this Epsilon opportunity.
(2) Comparisons with various types of instruments

Each of the tasks for investigating elemental-scale processes (Table 3-1 in NGSPM’s SOT report)
implies certain requirements for instrumental capabilities, such as spatial resolution, wavelength
range, temperature sensitivity, etc. The team identified notional ‘strawman’ instruments to execute
each observation. As a result, it becomes clear how many instruments are required for each subobjective. Table 3-1 in the NGSPM’s SOT report provides the list of key observations and notional
instruments (T-##). In consideration of a minimum set of instruments with which NGSPM canaddress the greatest number of high-priority Tasks, we find that a majority of the key required
measurements can be met with a suite of instruments including:
1.
2.
3.
4.
5.

0.3” coronal/TR spectrograph (T-09)
0.2”- 0.6” coronal imager (T-07)
0.1”- 0.3” chromospheric imager and magnetograph (T-04)
0.1” photospheric magnetograph (T-01)
0.1” chromospheric spectrograph (T-05)

While other combinations of candidate instruments would likely also produce significant results, thelist above represents the smallest set of instruments addressing as many as 32 of the 56 listed Tasks;
it is the list of instruments that the team feels should receive the highest priority for selection. Of
these 5 instruments, T-04, T-01, and T-05 can potentially be merged into a single telescope withfocal plane instruments, when suitable wavelengths in near IR, visible and near UV are chosen in
62/100the design. Therefore, in the subsequent sections, the team considered some mission concepts foraccommodating the following three instruments.
1. 0.3” coronal/TR spectrograph (T-09)
2. 0.2”-0.6” coronal imager (T-07)
3. 0.1” chromospheric/photospheric magnetograph/spectrograph (T-01/04/05)
The team recommended that NGSPM aims should be realized with a single-platform, as a JAXA
Strategic Large mission with contributions from NASA, ESA, and ESA member states. If thesingle-platform approach is not possible or available, a combination of two or three spacecraft canachieve many of the NGSPM objectives, with some loss of capability and at increased risk. Table
14.1 summarizes the results of the trade-off study among the nominated 3 + 1 instruments andSolar-C for JAXA’s strategic large mission with some additional aspects for trade-off, i.e., cost,size/weight, capabilities of domestic and international communities and data volume. As the result,
we reached the conclusion that the 0.3” coronal/TR spectrograph (T-09), i.e., EUVST is the mostpromising choice.
Table 14.1: Results of a trade study for selection of instruments

Science return,achievement levelof scientificobjectives
Uniqueness /synergy withother telescopes
Cost (fit toepsilon)Size/weight(fit to epsilon)
Technical andprogramatic risks
capabilities ofdomestic andinternational
communities
data volume, data
rate, and size ofpost analysis
Evaluation
summary0.3” coronal/TR
spectrograph(T-09 ~ EUVST)

0.2”-0.6”
coronal imager(T-07)A

B

A
Significant
advantages fromothers
A

B

A

A

A
No serious risksA
No serious risks

A
World-wide
support

B
Limitedinterests in
Japan
AHigh buttreatable

A
High but treatable
7A

A

4A, 3B

0.1”
chrom./photo.
magnetograph/spe
ctrograph(T-01/04/05).
A0.2”-0.5”
photo./chrom.
magnetograph(T-02)

B
some overlaps w/
4m groundtelescopeFhighF
Too large/heavy
B
Large and highprecision optics
B
Require large
resourcesC
moderate jumpfrom
Hinode/SOT
B
high
A

A
High but treatable

A
High buttreatable

2A, 3B, 2F

BA
No serious risksA

4A, 2B, 1C

Solar-Clarge mission(T-09+
T-01/04/05+
T-07)
A+

A+
Coordinatedobservations
F
Very highF
Too large/heavy
B
Large and highprecision optics
B
Require large
resources
A
High but treatable
2A+, 1A, 2B, 2F

The meaning of ranks: A – Excellent, B – Good, C – Fair, F – Fail(3) Instruments for spectroscopy

UV spectroscopy, which the 0.3” coronal/TR spectrograph (T-09) uses, was compared with soft Xray and visible-light spectroscopies scientifically discussed in section 6.1. This trade study was
made from various points of view, such as cost and risks, and summarized in Table 14.2. The study
shows that the UV spectroscopy is the most feasible method for performing the coronal
spectroscopic observations.

63/100Table 14.2: Results of a trade study for spectroscopic methods
UV spectroscopy(=EUVST)

Science return:Achievement level ofscientific objectives

A: HighNGSPM-SOT report

A: Identified as
key instrument
A: Treatable
A

Cost (fit to epsilon)
Size/weight(fit to epsilon)
Technical andprogramatic risks
Capabilities of domestic
and international
communities
Feasibility of key
Technologies
Availability of heritages
data volume, data rate,and size of post analysis
Evaluation summary

A: No serious risks
A
World-wide
supportA
See Section 12
TVLS grating
A: SoHO/SUMERSoHO/CDS
Hinode/EISAHigh but treatable
9A

Soft X-ray
spectroscopy(high dispersion)
B: Low(temperature
coverage)

B: Not identified as
key instrument
B: HighF
Too large/heavy
B: Big coolingsystem, long optics
B
Limited interests inJapan
BGrazing incidence
Segmented mirrors
Microcalorimeter
C: noneAHigh
1A, 6B, 1C, 1F

The meaning of ranks: A – Excellent, B – Good, C – Fair, F – Fail

Soft X-ray
spectroscopy(low dispersion)
B: Low(temperature
coverage, velocity
measurement)
B: Not identified as
key instrument
A: TreatableA
A: No serious risks
AInterests in Japan(PhoNEiX)
B
Grazing incidence
Segmented mirrors

Visible light
spectroscopy
B: Low (off-limb
observations,temperature coverage)
B: Not identified as
key instrument
A: Treatable
C
Large
B: Large & highprecision optics
B
Limited interests inJapan
B
Mirror coronagraph

C: None(Hinode/XRT)

B: SoHO/LASCO

A
High
5A, 3B, 1C

A
High
2A, 6B, 1C(4) Spacecraft and fairing accommodations for UV spectroscopy

Instrument designs of various lengths were considered. We investigated the spacecraft layout toaccommodate each of these designs in the Epsilon rocket nose fairing volume. This trade off study
determines the maximum length of the instrument. Three spacecraft configurations (Figure 14.1)
were compared with each other:
Case (1) The instrument is mounted on the +Z surface of the ISAS small standard bus.
Case (2) The instrument is obliquely mounted at one side (one surface from ±X and ±Y surfaces) of
the ISAS small standard bus.
Case (3) The instrument is located at the center, the lower portion of which is surrounded by thehollow-like bus.
The results from this trade-off study are summarized in Table 12.3, giving a conclusion that thelayout of Case (1) is more suitable, when the length of the telescope is about 3 m or shorter.
Case (1) is the layout in which the instrument is mounted on the +Z surface of the ISAS small
standard bus. This layout is fully based on the design standard of the ISAS small standard bus, inwhich mission payloads are basically mounted via interface fasters on the +Z surface of the bus.
The instrument is vertically oriented, and thus the maximum allowable length for the instrument islimited to about 3-3.5 m, depending on the interference with the taper shaper of the nose fairing.
The most important advantage of this layout is that the bus design is already verified, resulting in
the cheapest and less risk option for the spacecraft bus.
Case (2) is the layout in which the instrument is obliquely mounted at one side (one surface from±X and ±Y surfaces) of the ISAS small standard bus. Originally, case (2) was considered as one
layout option, in which the instrument is mounted at one side (one surface from ±X and ±Y
64/100surfaces) of the ISAS small standard bus. This may be similar to the mount configuration of EIS
and XRT on the Hinode spacecraft. However, the length of the instrument is limited due to the taper
shape of the fairing and the center of gravity has a fairly large offset from the axis of the rocket. It isdifficult to find merits in this layout. Thus, as a modified version, the instrument is obliquely
mounted at one side (one surface from ±X and ±Y surfaces) of the ISAS small standard bus. The
optical axis of the instrument is 10 degrees tilted from the rocket axis. This layout can reduce theoffset in the center of gravity. Moreover, it allows us to have the instrument with a length longer
than 4 m. The instrument is mounted at three points to the bus structure. The bus structure ismodified from the original small standard bus, giving additional cost and development risks.
Case (3) is defined by considering a layout similar to Suzaku and Hitomi. The bus structure isoctagonal prismatic and hollow shaped. The lower portion of a long-length instrument is surroundedby the octagonal prismatic bus structure. The heritage of the structural design from Suzaku andHitomi can be utilized, but this design increases the cost and development risks. Since the lower
portion of the instrument is fully surrounded by the bus, the heat dump at the radiators located at thelower portion, needs to be newly designed. For this purpose, heat pipes may be required from theheat source in the instrument to a radiator on the bus panel.
�������

�������

������� ��������

Figure 14.1: Accommodation of a 3.2m length telescope (40cm x 70cm x 320cm) to spacecraft
65/100Table 14.3: A trade-off study on the accommodation of a long telescope
Past heritage
Mass meritMassproperties
Layout design
Telescope
lengthLocation ofradiators

Case (1): +Z mount
A: Hisaki, Arase
B: Heavier because the standard bus
is used as is.

B: The center of gravity is located at
higher position. Lower rigidity.
A: The ISAS small standard bus
layout can be applied.
B: The telescope length is limited at
3-3.5m.
A: More flexible location for radiators in the instrument.

Alignment
design

A: The telescope axis can be aligned
to the rocket axis.

Development
Cost

A: The bus is already verified.
A: Cheapest because the small
standard bus is used as it.
A: Less risks for the bus because thesmall standard bus is used as it.
7A, 3B

Development
Risk
Evaluation
summary

Case (2): oblique side mountCase (3): hollow mount
A: Hisaki, Arase, (Hinode)
A: Hitomi, Suzaku
A: Can be optimized by
A: Can be optimized with areallocating components onhollow-type bus newly
the mount side.
developed.
A: The center of gravity is located at the position slightly lower
than case 1. Higher rigidity.
B: A large offset in center of
B: A new design based on
gravity from the rocket axis.
Hitomi bus.
A: 4m or longer telescope can be accommodated.
B: The lower portion is fully
surrounded by the bus. Needheat pipes.
A: The telescope axis can bealigned to the rocket axis.

B: The telescope axis is tilted
from the rocket axis, may give
demerit.
B: New verification of the bus is required.
B: Additional cost to modify a B: Additional cost for thepart of the bus.
new design.
B: No significant risks for the bus, but requires MTM/TTM.
5A, 5B5A, 5BThe meaning of ranks: A – Excellent, B – Good, C – Fair, F – Fail

(5) Survey of optical parameters

A survey of optical parameters was performed to determine the strawman design under following
constraints:•
Minimizing the number of optical elements in the
optical configuration should be taken to achieve
high temporal resolution, i.e., high throughput
performance. That is the optical configurationconsisting of the primary mirror, slit, grating, anddetectors.
•
Spatial sampling should be less than 0.3” perpixel at the CCD/IAPS to achieve a high spatial
resolution.
•
In order to fit the instrument into the payload
fairing of the Epsilon rocket, distances between
the primary mirror and the grating, and between
the grating and CCD/IAPS, should be equal to orless than 3100 mm.
•
In order to relax the curvature of the grating, thegrating-CCD/IAPS length should be as long as
possible, while the magnification of the grating
should be as small as possible, which makes thedevelopment of the TVLS grating more feasible.
Figure 14.2: Examples of the parameter•F/# (focal ratio) should be as large as possible to space that satisfy the conditions. Darkerachieve high resolution in a wider field of view.
colored regions show the larger number•
The diameter of the primary mirror should be as of solutions.
large as possible to provide a large effective area.
•
The optical parameters of LW bands have longer distance between the grating and IAPS,
and have more tight constraints than those of SW bands. We first determine the parameters
66/100of LW, and determine such parameters of SW that all bands have the same pixel size.
With the optical configuration, optical parameters are related as the following relations:
・ [F/# ( )] = [focal length (mm)]/[mirror diameter (mm)]・ [mirror-grating distance (mm)] = [mirror focal length (mm)] + [slit-grating distance (mm)]・ [spatial resolution at CCD/IAPS (”/pix)] = [pixel size at CCD/IAPS (µm/pix)]/([plate scale
at slit(µm/”)]<x[grating magnification ( )])・ [slit-grating distance (mm)] = [grating-CCD/IASP distance (mm)]/[grating magnification ( )]・ [plate scale at the slit (µm/”)] = [mirror focal length (mm)] x tan(1”) x 1000
・ [pixel size of IAPS in LW channel] = 20 µm (equivalent pixel size at the intensifiers’
surfaces)・ [pixel size of CCD in SW channel] = 13.5 µm
To find preferable parameters, we set the following inequalities:
・ [mirror-grating distance] ≤ 3100 mm
・ [grating-CCD/IAPS distance] < 3100 mm
・ [spatial sampling at CCD/IAPS] < 0.3”/pixel
・ [Grating magnification] < 8
The parameter spaces satisfying the inequalities are shown in Figure 14.2. Among the parameter
spaces, one of the solutions was chosen to gain the “largest mirror diameter”, “largest gratingCCD/IAPS distance”, and “largest F-number”. The optical parameters are listed in Table 14.4 andused as the baseline in section 15.
Table 14.4: optical parameters for the strawman design
Spatial sampling at CCD
Plate scale at slit
Mirror diameter
F/#
Grating magnification

0.184 ”/pix
13.57 µm/”
280 mm10
5.5 (SW), 8 (LW)

Focal length of the primary mirror
Slit-grating distance
Mirror-grating distance
Grating-SW CCD distance
Grating-LW IAPS distance

2800 mm
300 mm
3100 mm
1650 mm
2400 mm(6) Orbit and telemetry trade-off study

Trade-off studies on the orbit were performed for several orbit candidates to identify concerns thatneed to be dealt with. As a result of those trade-off studies (Table 14.5), the baseline orbit was
identified: a sun-synchronous polar orbit (SSO), similar to the Hinode orbit. This orbit wascompared with low earth orbit (LEO) with an inclination (31 deg), geo-synchronous orbit (GSO)
with an inclination (28.5 deg) and a Halo orbit at the Lagrangian point L1. The LEO was not an
option because of the difficulty of the thermal design and the continuity of observations. The GSO
and L1 orbit were not an option from the point of view of the difficulty of carrying a 400-500 kg
class spacecraft by an Epsilon vehicle.
The following required items were primarily studied for each orbit candidate:a. High-rate science telemetry downlink
The orbit will allow downlink of the mission data continuously produced with the average rateof data output specified in section 6.4.
b. High pointing stability
Stable thermal environment with less orbital variation.
c. Thermal design
The orbit shall ensure that the thermal design of the EUVST instrument is technically feasible.
d. Continuity of observation
Shorter eclipse seasons, shorter duration of each eclipse and minimal orbit maintenance ispreferable.
67/100e. Real-time science operations
Real-time operations are carried out to upload the science plans to the spacecraft. Some science
data downlinked are used to make fine pointing adjustment.
f. Feasibility of the launch by Epsilon rocket.
Table 14.5: Result from orbit and telemetry trade-off study
Low earth orbit (LEO)

Sun-synchronous polar
orbit (SSO)

Altitude
Inclination
PeriodHeritage(solar obs.)
High-ratesciencetelemetry
downlink

500-700 km
500-700 km
31 deg
97-98 deg
98 min98 minHinotori, Yohkoh,
Hinode, TRACE, IRISRHESSI
B: X-band (8Mbps) is recommended and requires>15 passes (10min each) per day. Supports by
ESA and NASA stations are essential to increase
the telemetry.

Highpointing
stability

A: No antenna pointing mechanism is needed,giving no micro-vibration.

ThermaldesignF: day/night transition
every orbit causes large
thermal deformation,
giving difficulty inhigh resolution.

Continuityofobservation

F: The longest duration
is 60 minutes. A night(~30 minutes) every
orbit.

Real-timescience
operation

A: Real-time access ispossible but it isrestricted to 10 minutes
each at 5 USC passes.

Feasibilityof the launch
by Epsilon
rocketEvaluation
summary

A: No critical issue

3A, 1B, 2F

B: The orbital variationby the periodic infrared
radiation is a concern
for high resolution
performance. Earthalbedo & IR radiation
increases the size of theradiators required.
A: The longest durationis 8 months. An eclipse
season (for 4 months)
with about 20 minutes(at maximum).

A: Real-time access ispossible but it isrestricted to 10 minutes
every 98 minutes for apolar ground station.

4A, 2B

Inclinedgeosynchronous orbit(GSO) 1)
36,000 km
28.5 deg
1 day
SDO

Halo orbit at theLagrangian point L1
L1 point
SoHO

A: X-band is
B: X-band (8Mbps) is
recommended andpossible with 30W
required ~3hrs contact
transmitter. Needs ~3
with USC 34m or
hr occupation of the
Katsuura.
64m Usuda antenna.
B: The motion speed of B: The motion speed
the antenna is slowof the antenna is slow(~30 deg/hr), but need(~15 deg/month), but
to evaluate the microneed to evaluate thevibration level.
micro-vibration level.
A: The orbital variation is very small, providing
a stable thermal environ- ment. Earth albedo
and infrared radiation is very small, making thethermal design easier.

A: Eclipse seasons
A: No eclipse season.
with 70 minutes (atOrbit maintenancemax) per day, eachmaneuvers (several
continues 20 days,
times a year) aretwice every year.
needed every year.
Related potential
demerit is to need alarge capacity of thebattery. Orbit maintenance maneuvers (twicea year) are needed.
A: Real-time access can be arranged in the
duration of pass (8 hr maximum).

F: The vehicle cannot carry 500kg class mass tothis orbit
4A, 1B, 1F

3A, 2B, 1F

The meaning of ranks: A – Excellent, B – Good, C – Fair, F – Fail
Note 1) GEO (Geo-stationary earth orbit, inclination 0 deg) is almost equivalent to GSO except for the orbit
inclination, and brings few benefits compared with GSO. Rather, it requires more ~200 kg propellant toinstall the spacecraft to the GEO, compared with inclined GSO.

68/10015.

Baseline architecture (draft)
ベースラインアーキテクチャ (ドラフト)
15.1. Subsystems and their dependencies [12-3-2]
サブシステムとその間の要求の依存関係
15.1.1. EUVST Instrument Module Design

The EUVST instrument consists of two portions, i.e., a Telescope Unit and a Spectrograph Unit
(Figure 15.1). The entire package will be mounted to a mechanical interface at the top panel of thebus module. This configuration needs to provide thermal independence from the spacecraft. The
Telescope Unit structure is the main support structure of the instrument module and has provisions
to mount the Spectrograph Unit structure. The opto-mechanical layout of the system is shown in
Figure 15.2. The Telescope Unit will consist of a front door assembly, the single-mirror telescopeassembly (an off-axis parabola with pointing capabilities), the guide telescope, the telescopeelectronics, the heat rejection assembly and the deflector plate assembly. It provides a solar image
in the focal plane, the entrance to the Spectrograph Unit, and provides image stabilization and slit
scanning capabilities. The Spectrograph Unit houses the slit assembly, the slit imaging assembly,
the grating assembly, the focal plane assemblies, CCD electronics box, the intensified APS (IAPS)
electronics box and the detector radiator assembly. It accepts the light from the telescope passing
through the entrance slit towards the grating and the detector assemblies.
spectrograph unit

VUV IAPS
assembly
EUV CCD
assembly

grating
assemblyslit assembly

telescope unit
Sun

IAPS electronics box

slit imaging
assembly

CCD electronics
box
thermisters/heaters

Sun

TC & TM
from S/C

heat rejection ass.
deflector
plate
assembly

aperture
door
assembly

SpectrographElectronics
Box

Power
from S/C

primary
mirrorassembly
guide telescope assembly

telescopeelectronics
box

thermisters/heaters

Figure 15.1: Physical block diagram of the instrument, showing major components in the
instrument and digital communication links.

Figure 15.2: The instrument opto-mechanical layout
69/10015.1.2. EUVST Optical Configuration and Assembly Dependencies

The optical configuration follows the two-element design pioneered by the Hinode/EIS instrument.
This two-element design minimizes the number of reflections in the system – an essential feature in
VUV instrumentation. Minimizing the number of reflections in the system is essential to meet thescientific requirement for high throughput. A schematic view showing the layout of the opticalelements is given in Figure 15.3. Characteristics of the telescope are given in Table 14.4.

LW IAPS detector

Figure 15.3: EUVST optical schematic. The figure shows the nominal geometry of the EUVST
design. The path of EUV radiation through the instrument is shown.
The slit selects a one-dimensional portion of the solar image and this radiation is incident onto a
concave diffraction grating. The beam from the slit is incident on the SW and LW gratings. These
two grating halves are mounted in close proximity. The gratings are figured and coated to optimize
the image quality and efficiency of each pass-band. The highly corrected off-axis toroidal gratings
with variable line spacing provides excellent image quality at the focal plane across all wavelength
ranges. Table 15.1 describes the grating geometry and the spectrograph characteristics. The imaging
performance of the combined telescope and SW/LW spectrograph at the central wavelength of each
CCD/IAPS plane at the slit center, ±50, and ±150 arcsec from the slit center is shown in Figure15.4. The instrument performance improvements are the result of a remarkable breakthrough inspectrograph design which incorporates a Toroidal Variable Line Space (TVLS) grating to performthe necessary measurements on the critical spatial and temporal scales. The highly magnifying
mount allows an instrument of reduced size compared to that of a conventional spectrometer withsimilar spatial resolution.
Table 15.1: Spectrograph design characteristics

LW wavebands
1st order: 690-850 Å, 925-1085 Å, 1115-1275 Å
2nd order: 463-542 Å, 557-637 Å
Detectors
Intensified APS, two 1920 × 2048 devices for a
4096 × 2048 CCD, 13.5 µm pixels
total of 3840 × 20481), 20 µm pixels
Geometry and4200 grooves/mm2000 grooves/mm
dispersion
165 cm grating to detector
240 cm grating to detector
20 mÅ/pixel
37 mÅ/pixel
Plate scale
0.184”/13.5 µm pixel
0.184”/20 µm pixel
Slit length368” (2000 pixels)
368” (2000 pixels)
Magnification
5.5
8.0
1) The 2048 will run along the spatial direction and 3840 along the spectral domain.
Spectral ranges

SW waveband
170-215 Å

70/100Figure 15.4:
Imaging
performance of the
combined telescopeand SW/LW
spectrograph at 18.5,
74.3, 100, and 122
nm. Spots shown arefor slit center, ±50,
and ±150 arcsec
from the slit center.
FWHM of the spot
size in μm is shownin the left bottom ofeach panel.

15.1.3. Slit Imaging Assembly

The Slit Jaw Imaging system uses relay optics to form a solar image including the entire slit and its
surroundings on an APS sensor, through relatively narrow bandpass spectral filters to isolate
specific wavelength ranges. The concept design demands a FOV of a few hundred arcseconds, withspatial resolution of ~0.5 arcsec and pixels of ~0.16 arcsec; the final design requires a more detailed
definition of the slit jaw itself and the envelope and structure available for mounting components.
The camera will be a simple framing camera with flight heritage; there are no challenging
requirements for speed, read noise or operating modes. A rolling curtain shutter is used to vary
exposure time, from tens of milliseconds to seconds as needed, with optional automatic exposure
control to avoid saturation during flares. Structural/thermal/optical analysis of the flight structure
will determine whether a focus mechanism is needed; IRIS did not have independent focus for its
slit jaw system. The baseline slit jaw imaging system is a 1600 Å broadband system, similar to thaton SDO/AIA (Lemen et al, 2012). This would use one or more interference filters on MgF2
substrates to reject the visible light and isolate a band that includes strong lines of He I, Fe II and C
IV. The 1600 Å region generally shows the upper photosphere but includes the transition region
emission during flares. A filter wheel could be used to provide additional channels such as the 1700
Å and white light channels of SDO/AIA.

A more interesting possibility would be to follow the scheme used by the IRIS slit jaw system, thatprovide narrowband images of the chromospheric Mg II k line emission and nearby bands ofcontinuum and photospheric lines. The Mg II images show chromospheric structures totally absent
from the 1600 or 1700 Å images, which have scientific interest themselves and would provide amore accurate co-alignment with DKIST chromospheric images or MUSE transition region images.
The key component of this system is a Solc birefringent filter with a FWHM of 3.6 Å and free
spectral range of 33 Å (Berger et al., 2012). This filter contains no moving parts or oil; the opticalelements are mated with high-viscosity optical grease, as the Lyot filters flying on SoHO/MDI andSDO/HMI, and are hermetically sealed in an enclosure. A filter wheel with blocking filters of ~15 Åbandwidth centered on 2796 and 2832 Å provides the Mg II k images and far wing images, whichshows photospheric granulation and bright points. A four-position filter wheel could have anadditional clear filter for ground testing and another UV blocking filter, perhaps on the strong Fe II
lines at 2630 Å for upper photospheric images.
71/10015.1.4. Image motion compensation and raster system (IMC)

To obtain the required observations, the telescope line of sight must be precisely pointed to the Sun.
During a single exposure (typical 1 s), the line-of-sight must be maintained with an error of< 0.2”(3σ) to obtain an overall resolution of <0.4” (Table 15.4). Moreover, the solar image focusedon the slit must be moved to make mapping in a field of view. An image motion compensation andraster system takes care of these functions. As shown in Figure 15.5, the combined capability of the
tip-tilt mirror and the guide telescope will allow image motion stabilization and raster scanning. An
externally mounted guide telescope provides the necessary aspect information by measuring theposition of the solar limb with four redundant photodiodes. Intensity differences between opposite
diodes are used to derive a displacement error signal. A fast steering mirror provides the mirror
tip/tilt articulation. Similar to IRIS, raster steps are implemented as offsets within the main control
loop. The system will operate at >10Hz frequency and will provide the required 0.2”(3s)stabilization. Line of sight jitter with frequencies above the effective bandwidth will be addressedby the control of micro-vibrations (Section 12.2.5). This is an open-loop control system in the sensethat the error signal is derived from the GT rather than from motions of the image in the telescope
focal plane. Therefore, careful calibration is needed to match the voltage scale of the tip/tilt
actuators with that of the GT error signals, and this is routinely achieved to accuracy on IRIS of <2%. A pair of independently rotatable optical wedges (Risley prisms) are mounted between theentrance filter and the objective lens so that the GT can be oriented off the optical axis of the
primary mirror by up to 21 arcmin. When the ACS repoints to a target away from disk center using
the UFSS, the wedges are rotated so the image will be approximately centered on the GT sensor,
within its linear range. Table 15.2 details the expected performance of the mirror assembly, theguide telescope and the servo-control loop.

Figure 15.5: Image motion compensation and raster system functional block diagram.
Table 15.2: Guide telescope and tip-tilt characteristics

Guide Telescope:
Tip-tilt Mirror:
Characteristics and Performance based on IRIS Servo Control Loop Operating Characteristics
Wavelength
Overall operation> 50Hz nominal
l0 = 570nm, dl= 50nm
frequency
Focal length
1.88 mAcquisition range1)
Sensor sampling
250 Hz> ± 21 arcmin
frequencyLinear range> ± 80 arcsec
Noise equivalent
0.02 arcsec rms
Expected bandwidth
>10Hz
angle
cutoff
Note 1) The line of sight can be adjusted with wedges.

15.1.5. Thermal design and interface requirements

The instrument shall be designed as a passive, cold biased thermal system, optimized forinsusceptibility to external heat sources. It draws strong heritage from the spectrometers on SoHO
and on Hinode, where experience has shown that an ultra-stable temperature equilibrium of the
72/100optical bench is required to reach high optical performance characteristics. Long-term variations(seasonal, aging effects) are less critical. These requirements put severe constraints on the thermomechanical S/C interface, and call for an isostatic mount of the instrument to prevent thermo-elastic
cross-effects and to minimize radiative and conductive coupling to the S/C. The sun synchronous
orbit (SSO) provides a relatively stable thermal environment. The varying effects of Earthshine andAlbedo on the instrument are minimized by optimal radiator placement and orientation. In SSO, aneclipse of up to 20 min will occur every orbit for 3-4 months in a year. Compensation heaters will
minimize the post-eclipse recovery time.
An analysis based on a simplified 16-node thermal math model has shown that the thermal concept
as shown in Figure 15.6 can be separated into three basically independent blocks.

Figure 15.6: Schematic drawing of the thermal design and results obtained from a 16-node thermal
model to verify the conceptual design.
1) Telescope mirror and open-aperture section: With an absorption coefficient of 0.7 for its surface,
the mirror will absorb most of the solar input. An alternative coating with reduced alpha will be
studied during Phase A as described in section 16.3. The mirror will be cooled passively with an
external radiator as shown in Figure 15.6. Optical performance requirements do not allow
conductive cooling. The mirror and mount will be designed to operate at elevated temperatures
while maintaining the optical figure. The mount design will follow the successful example ofHinode/SOT which was designed for temperatures up to 47°C. A very similar design was used for
SUMER (Wilhelm et al., 1995) and is currently implemented for the mirror of the SPICE
spectrometer of Solar Orbiter (Schühle et al., 2007).
2) Pre-slit heat rejector: The pre-slit, which collects most of the light reflected from the telescope
mirror, will protect the slit from excessive heat load by coupling to a second radiator in thetelescope. The section behind the spectrometer slit is regarded as free from solar heat input.
3) Focal Plane Assemblies: The CCDs and APSs require mechanical stability, thermo-elastic
insusceptibility, cooling (<-60°C for the CCDs and <-45°C for the APS) and optimized
accommodation of the proximity electronics. These requirements will be achieved by an advanced
design that is based on Hinode/EIS experience. The SW CCDs will be thermally isolated from thestructure and cooled using a flexible strap to an external radiator. Depending on the results of the73/100Phase A modelling, the EUVST team will incorporate an active thermoelectric cooler if required.
The LW IAPS will be cooled using an additional external radiator.
15.1.6. Electrical System and Interfaces
The electronics block diagram is given in Figure 15.7. The power and digital interface with thespacecraft is through the spectrograph electronics box. As shown in Fig. 15.7, the spectrographelectronics box (SEB) receives uplinked commands from the spacecraft and communicates directlywith the transponder. The SEB power converter provides conditioned power to the SEB box. The
SEB relay card provides switched power to the various detector electronics boxes and the telescopeelectronics box. Communication within the SEB is with a high speed backplane. A spacewireinterface card services the various EUVST spacewire links. The SEB incorporates a mass memorystorage card with 39Gb. This provides a buffer equivalent to roughly 12 hours of downlinked
telemetry. A mechanism driver card services the slit exchange mechanism, the grating focus
mechanism, the shutter mechanism for the EUV CCD, the filter wheels for the slit jaw imaging
system and the order sorting filter, three reclosable detector doors and the guide telescope.
Thermister, heater and thermoelectric coolers are controlled with a separate card. A microprocessor
card provides ample capability to control and sequence spectrograph operation. During phase A,
additional possibilities will be evaluated. The SEB box architecture is similar to the Secchi
electronics box provided by NRL to the STEREO spacecraft.

Figure 15.7: Block diagram of Electronics in EUVST
74/100Separate electronics boxes are envisioned for control of the CCD and IAPS detectors as shown in
Figure 15.7. The detector electronics will also provide compression of the spectra and formatting of
the data into CCSDS packets. Compression will be implemented with a CCSDS 122.0 compatible
compression chip (for example, CWICOM). The spectra will be compressed using both lossy(JPEG2000, wavelet compression and H-compress) as well as lossless (for example, RICE andDPCM) compression. The JPEG compression is currently being used on EIS and has a compression
factor of 4-5. The software will also have the capability to bin pixels as well as provide overall
image statistics. An analysis shows that the CWICOM chip has sufficient capacity(60Mpixels/second) to process the output of two ADCs (4Mpixels/second total) in each camera
head. Sustained operations of observing modes with a telemetry rate of >0.8Mbps would require
lossy compression (~8) to fit into the available downlink. The front end electronics of the detector
electronics boxes will be nearly identical to electronics provided for SO/Metis, SO/SoloHI,
SO/EUI, STEREO/Secchi and Hinode/EIS. The high voltage power supplies are expected to be
identical to supplies provided by MPS for SO/Metis and SO/EUI. The APS slit jaw camera
electronics will be a reduced functionality version of the IAPS camera.
The telescope electronics box will provide control of the tip/tilt mechanism, the focus mechanism,
the door and deflector plate voltage. Mirror voice coil actuators will be operated using customized
analog/digital drivers; direct feedback of the position will be available from the capacitancemicrometers. Operations of the telescope electronics box will be controlled by an embedded
microcontroller.
15.1.7. Spacecraft system

The system block diagram is shown in Figure 15.8. Our proposed baseline for the spacecraft system
is to utilize the bus system defined in the ISAS small standard bus, as also shown in Table 15.3.
However, as discussed in section 10.2, a few modifications are needed to meet the high pointing
stability and high mission telemetry.
Mission payload
SpectrographE-box

memory included
Mission data
output

CCD elec box
IAPS elec box
Telescope E-box

SWR3
TCU

DR

Figure 15.8: System block diagram.

75/100HK data(system, mission)

Table 15.3: List of components in the proposed baseline. Highlighted items are not available on thestandard bus.
Component

MissionMission payload
EUVST
X-band telemetry system

Code

#

Used in proposedbaselineEUVST
XMOD, XPA, XFIL
X-SW, XANT
TCU

1
1 set

√
√

1

√

1
1
1
2

√
√
√
√

√
√
√
√

3
2
1
1
2

√
√
√
√
√

√
√
√
√
√

2
0
1
1
2
1

√

√

√
√
√
√

√
√
√
√

2
4
3
1
1
2
0
1
1
6~7
1 set

√
√
√
√
√ (upgrade)
√

√
√
√
√
√
√
√

1

√

Telemetry & command unit
Bus system
Management (SMS)
Satellite management unit
SMU
Data recorderDR
Telemetry/command IF module
TCIM
Spacewire router
SWR
Radio Communication (RF)S-band antenna
SANT
S-band diplexa
SDIP
S-band switch
SSW
S-band hybrid
SHYB
S-band transponder
STRP
Power Supply (EPS)
Solar array paddle
SAP
Paddle drive motor
SADM
Power control unit
PCU
Array power regulator
APR
SAP blocking diode
SBD
Battery (50Ah)
BAT
Attitude and orbit control system (AOCS)
Attitude orbit control computer
AOCP
Reaction wheel assembly
RWA
Magnetic torquer
MTQStar tracker
STT
Inertia reference unit
IRU
Coarse sun sensor
CSAS
Sun presence sensor
SPSH
Ultra fine sun sensor
UFSS
Geo-magnetic sensor
GAS
AOCS interface moduleAC***
Thruster system (RCS)
RCS
Thermal control (TCS)
Heater control electronics
HCE

√
√
√
√

Ref: Default
standard bus

√
√
√
√

For the high pointing stability, Ultra Fine Sun Sensor (UFSS, sections 12.2.3, 16.2, and 16.3) will
be used instead of sun presence sensor as a sensor for measuring the accurate direction of the Sun inreal time and delivering the measured angles to the attitude control computer for controlling theattitude. Gyroscope (Inertia Reference Unit; IRU) will be upgraded from the standard one to higher
resolution IRU (for example, pulse wait 0.05” to 0.005”). Higher resolution IRU means a narrower
range of measurements (for example, 4 deg/sec to 0.4 deg/sec). Additional studies associated withthis modification will be needed during the design phase, such as the acquisition of the Sun after thelaunch and contingency cases with the given battery capacity. A thruster system is assumed
according to an initial study in section 15.4.2.
High telemetry mission data line is required in the mission. An X-band telemetry system with alarge mass data recorder will be accommodated for this requirement. Since no space is available in
the ISAS small standard bus, a small electrical panel structure for accommodating the electricalcomponents consisting of the X-band telemetry system and data recorder will be added andmounted on the +Z surface of the bus.

76/10015.2. Risk mitigation [7-3-2]
リスクの mitigation
The major risks of the proposed mission concept were listed and briefly discussed in section 12.4.
Four major risks have been identified and all of them are related to the requirement for spatial
resolution. For the requirement for spatial resolution, we define a strawman budget table as shownin section 15.3, in which the target value is given to each of error sources. The design andevaluation studies will be carried out to confirm whether each target value is achievable. Depending
on results from the studies, the budget table will be updated. Tests for verification of the flightmodel will be carried out to confirm that each term in the budget table is achieved and finally thatthe flight model has the required performance. The major risks listed in section 12.4 are a part of theterms in the budget table. Here we explain how to proceed for the major risks and discussalternative considerations for mitigating the major risks.
• No. 001: The grating performance: The grating parameters have been determined for the baseline
design. For confirming the manufacturing feasibility, it is important to study the tolerance in theparameters and its manufacturing feasibility with grating providers. We have initial contacts witha few grating providers (Horiba JY and Zeiss). These contacts have shown that the grating
specifications are achievable. Further iteration on the grating design and its tolerance analysis
with the grating providers is planned during phase A. Prototype gratings will be procured andtested during Phase A as described in Section 16.3.4. We will also have plans for backup
solutions (spherical VLS grating, reduced groove density).
• No. 002: Wavefront performance of the primary mirror: The primary mitigation will be tocomplete the thermal-structure design and analysis for the baseline early in Phase A. The
analysis will provide the mirror deformation as a function of temperature and mechanicalmounting arrangement. Temperatures up to +80°C will be considered. The EUVST team will
also finalize the primary mirror coating during Phase A and measure the coating reflectivity inthe visible and UV. To avoid contamination of the primary mirror, the operating temperature of
the mirror will be >50°C.
• No. 004: Micro-vibration from momentum wheels: A first micro-vibration analysis with the
ISAS small bus model carried out in summer 2017 (section 12.4 and 16.2.5) indicated that the
micro-vibration from momentum wheels have potential to exceed the 0.1 arcsec (3s) level,giving blurring to the spatial resolution at unacceptable level. It is important to define moreclearly the direction of design before going to phase A study. A large uncertainty exists in thefirst micro-vibration analysis, efforts for mitigate it focus on reducing the magnitude of theuncertainty, by 1) carrying out an experimental measurement of the transfer function with the
ISAS small bus as well as 2) by starting to discuss the amount of micro-vibration with wheel
vendors during pre-Phase A study. These two efforts will improve our model fidelity and thesecond micro-vibration analysis with the updated model would provide more accurate prediction
on the micro-vibration level. If the results from the improved micro-vibration analysis still showthat the vibration level exceeds the allocated value, we will take the following actions: 1) the
trade-off studies of the isolation damper (with the major risk numbered No. 012), and 2) therevision of the budget table. If those actions do not allow to meet the spatial resolution
requirement, we shall consider a slight relaxation of the requirement.
15.3. Worst case analysis [8-3-2]
最悪ケース解析High spatial resolution is one of the key challenges of the EUVST. The EUVST team has assessed
the image quality of the baseline design and created an imaging error budget. The imaging error
budget contains the major error components as well as significant margin. The results are
summarized in Table 15.4.
77/100Table 15.4: Initial budget allocation for achieving the FWHM spatial resolution requirement.
Term
Optical design

FWHM1
0.11”
0.19”

Remarks
SW Avg Spot size: 6µm in FOV ±50”
LW Avg Spot size: 15µm in FOV ±50”
Based on Hinode/EIS grating experience

Spectrograph fabrication and
0.09”
assembly
Telescope diffraction-limited
0.16”
13nm rms wavefront error is assumed;
Performance at 182nm2
Scaled performance of SOT primary mirror
Total optical resolution
0.22”
0.27”
Pointing jitter
0.16”
Table 10.1: <0.2” (3s)Total EUVST Resolution
0.27”
0.31”
Margin vs 0.4” 3
50%
29%
Table 5.1: Science requirement
1) FWHM: SW (left) and LW (right)2) If the telescope has a 26nm rms wavefront error, the total EUVST resolution is 0.4” for the SW case.
3) EUVST spatial resolution performance has significant margin which accommodates a reasonable rangeof worst case scenarios. The threshold requirement of 0.8” FWHM is not expected to be reached during
the mission development.

The term “optical design” is the optical performance of the entire instrument that has been
confirmed with ray tracing, as shown in Figure 15.4. Thus, this is the worst case for the opticaldesign and we will continue to optimize the optical design for more better spatial resolution. 13nm
rms wavefront error which is diffraction-limited wavefront error for 182 nm is allocated to thetelescope as the initial budget. The active mirror assembly and the telescope structure formaintaining the positional relations among optical elements will be developed to achieve thisallocated value. This allocated wavefront error is about a half of the wavefront error achieved in the
primary-secondary mirror configuration of Hinode/SOT. The allocated error is roughly scaled from
Hinode/SOT, by considering the difference in mirror diameter (SOT 50cm vs. EUVST 28cm). Even
when the telescope has a 26nm rms wavefront error, the total EUVST resolution is 0.4” for the SW
channel. The SOT’s achieved wavefront error was designed in the temperature up to 47°C. Since
the EUVST mirror is operated at a high temperature (currently 80°C), thermal deformation may becaused in the temperature change from 47°C to 80°C and this is currently unknown quantitatively.
The mirror assembly design will be carried out to minimize this thermal deformation in pre-Phase A
and Phase A study. On the other hand, as described in section 16.2.3, the team is studying a coatingwith enhanced reflectivity and significantly lower solar absorption and this effort will be helpful to
reduce the mirror temperature significantly.
Assessment of Pointing Stability
With the pointing signals from UFSS, the pointing stability of ISAS standard bus was evaluated by
the bus company, based on knowledge from Hisaki and ASNARO-1. Table 15.5 shows that it would
be possible for the spacecraft systems to achieve the science requirements given in Table 10.1,
although further studies are needed in the co-alignment between UFSS and payload instrument,
UFSS internal variation, and micro-vibration from reaction wheels. Note that the IRU is upgradedto a higher accuracy component.
Table 15.5: Summary result from an initial analysis of the spacecraft pointing stability
Short (<10 [sec])
Medium (1 [hour])
Mission life
(absolute pointing)

Requirement
Evaluation result
Requirement
Evaluation result
Requirement
Evaluation result

qx [arcsec] (3s)
0.2
0.12
4.2
1.89
42.0
3.89

78/100qy [arcsec] (3s)
0.2
0.12
4.2
1.89
42.0
3.89

qz [arcsec] (3s)
7.00
0.57
212.0
14.53
212.0
117.52

The magnitude of spacecraft jitter in the short timescale (corresponding to individual
exposures) may exceed the budget in Table 15.4 and, thus, EUVST is required to have an image
motion compensation system in the instrument (section 15.1.4). Additional sources of error such as
drift between the UFSS and the telescope optical axis were not considered in the analysis. The noise
equivalent angle from the guide telescope is 0.02 arcsec rms (Table 15.2), which will be sufficientto achieve 0.2 arcsec (3s). On IRIS, the use of the GT error signals with the tip/tilt secondary mirror
keeps the solar image stably positioned on the slit with a precision of <0.05 arcsec. A similar
performance would meet the EUVST requirement jitter requirement in Table 15.4.
Optics Alignment Tolerance
The error budget for the entire telescope and spectrograph will be divided into terms, such as thedesign, fabrication errors of the optics, errors from mechanical and thermal deformation, and focus
positional error. We will update the budget table in the development efforts (section 16.3). We
should note that the wave front error of 13nm rms (equivalent to 0.16”) was allocated to achieve thediffraction limited performance. The Hinode/SOT experience indicates that 13nm rms is achievable.
The study of the mirror figure deformation is one of the key areas because of the high operationaltemperature.

Figure 15.9: Instrument alignment sensitivity. FWHM of the spot diagrams at the LW detectors as afunction of the distance between the mirror and the slit, the slit and the grating, and the grating andthe detectors. Each column shows the different wavelength. Colors display different incident angle
along the slit direction described in the left-top figure. The plate scale of the detectors is 20 µm per
pixel.
The optical performance computed by changing the distances between optical elements are shownin Figure 15.9. This figure shows that the adjustment of ±50 µm is needed for the distances between
the primary mirror and the slit, and between the slit and the grating to fulfill the spatial resolution of0.4” FWHM in the field of view of ±50”. The distance between the grating and the CCDs are, on
the other hand, adjusted within 5 mm. Assuming that the accuracy of the alignment during the
integration of the optical elements is about 100 µm, focus adjustment mechanisms are considered in
79/100the primary mirror and the grating as the baseline concept. Tolerances of the decenters of the
primary mirror, slit, grating, and CCDs are 500, 500, 1000, and 200 µm, respectively. This meansthat optical elements have no need of positional adjustment in non-focus directions.
15.4. Space environment [3-3-2]
宇宙環境
15.4.1. Lower limit in the altitude of sun synchronous orbit

A sun synchronous orbit is selected for the proposed mission to satisfy the requirement of theseamless observation for 24 hours per day for nearly nine months of the year. However, the line-ofsight atmospheric length of the Earth becomes longer in the SSO orbit than that in the smallinclination low-earth orbit (LEOs), which leads to a low atmospheric transmission in EUV and UV
wavelengths by absorption processes. The absorption is reduced when the orbital altitude is higher.
The lower limit of the altitude in the requirement is determined by the atmospheric absorption, andthe upper limit is given from the radiation environment to be allowed.
The absorbers are mainly molecular nitrogen, molecular oxygen, and atomic oxygen in thethermosphere, and the density of the absorbers varies according to the solar activity. The absorption
cross sections are large for wavelengths of 40-80 nm and largely fluctuate between 80 and 100 nm
depending on wavelength. The deep absorption occurs when the axis of the Earth is tilted to thedirection of the Sun. The nominal orbit altitude of 630 km in the Solar-B mission was determinedfrom the EUV observations at the longest wavelength of 29 nm by considering the atmosphericabsorption at the solar maximum in 1991.
0.5

Orbital Phase [deg]17.1nm (h=500km)

17.1nm (h=600km)

Orbital Phase [deg]17.1nm (h=650km)

17.1nm (h=700km)

300

300

300

200

200

200

200

100

100

100

100

0

0

0

0

100

200
DOY

300

0

100

200
DOY

300

0

46.5nm (h=600km)

100

200
DOY

300

0

300

300

200

200

200

200

100

100

100

100

0

0

0

100

200
DOY

300

0

100

200
DOY

300

0

63.0nm (h=600km)

100

200
DOY

300

0

300

300

200

200

200

200

100

100

100

100

0

0

0

200
DOY

300

0

77.0nm (h=500km)

100

200
DOY

300

0

77.0nm (h=600km)

100

200
DOY

300

0

300

300

200

200

200

200

100

100

100

100

0

0

0

200
DOY

300

0

100

200
DOY

300

0

100

200
DOY

300

300

100

200
DOY

300

100

200
DOY

300

200
DOY

300

77.0nm (h=700km)

300

100

0

77.0nm (h=650km)

300

0

200
DOY

63.0nm (h=700km)

300

100

0

63.0nm (h=650km)

300

0

100

46.5nm (h=700km)

300

0

0

46.5nm (h=650km)

300

63.0nm (h=500km)
Orbital Phase [deg]

1.0

300

46.5nm (h=500km)

Orbital Phase [deg]

0.6
0.7
0.8
Transmisstion

0

0

100

Figure 15.10: Atmospheric transmission of the Earth to several wavelengths to be observed with
EUVST for the spacecraft altitude of 500, 600, 650, and 700 km. Day of the year (DOY) is in thehorizontal coordinate, and the vertical coordinate is the phase angle of the spacecraft orbit aroundthe Earth.
80/100We have calculated the atmospheric absorption as a function of zenith angle of the Sun from thespacecraft view, solar and geomagnetic activity represented by the F10.7 radio flux and Ap-index
with the help of Hedin’s thermospheric model. The zenith angle of the Sun seen from the LEO
spacecraft changes within a short-duration spacecraft orbit around the Earth and in the orbital
motion of the Earth around the Sun. Figure 15.10 displays the atmospheric transmission at therelatively high solar activity (F10.7=150 and Ap-index = 25) as a function of the orbital phase of the
Earth in DOY (day of the year) and the spacecraft orbital phase in the horizontal and vertical axes,
respectively. The zero degree in the vertical axis is the ascending node of the spacecraft to the
Earth’s equator, and 90 degree is the north pole in the case that the launch date is near the autumn
equinox point like Solar-B. The eclipse of the Sun by the Earth occurs in the May-to-Sep season for
the LEO spacecraft in SSO. The period of non-eclipse and non-absorption conditions for all EUVST
observing wavelengths is not long for the altitude of 500 km. The lower-limit altitude of 600-700
km will be the solution and is achievable with the Epsilon.
15.4.2. Necessity of propulsion system

The stability of a sun synchronous orbit was studied to know the necessity of propulsion systemonboard. The requirement from the science is that the total duration of the non-eclipse season is >8months per year within two years after the launch (section 10.1.2). Two uncertainties wereconsidered: 1) The injection accuracy of Epsilon rockets (Table 4.1), and 2) the height reduction due
to air drag at 2024-2026. For 1), the total duration of the non-eclipse season is reduced to less than 5months in the 2nd year if the installed orbit is 20 km lower than the target height and +0.2 deglarger than the target inclination. For 2), we obtained that the height reduction is 60 km/year (170m/day), 18 km/year (50 m/day), 6 km/year (17 m/day) for the orbit height of 550 km (air density:
8.5x10-13 kg/m3), 650 km (2.5x10-13 kg/m3) and 750 km (8.4 x10-14 kg/m3), respectively. The heightchange caused by air drag will further increase the deviation from the sun synchronous condition.
This initial study showed that the propulsion system should be included onboard the spacecraft aspart of the baseline architecture.
15.4.3. Spacecraft environment

The EUVST has assessed the radiation, thermal and orbital environment as part of the proposal
effort. The expected orbit is nearly identical to Hinode/SOT and is relatively benign.
15.5. Planetary protection [14-3-1]
惑星検疫
Not applicable.
15.6. New ground system [6-3-5]
新規地上系
Additional ground system software for converting the observations envisioned by scientists into
data will be written. This software will be based on the heritage of Hinode/EIS, but will also include
new ideas from more recent missions, such as IRIS. The first element of the ground system will be
the software for generating and validating user-defined observing programs and storing them in a
database. The next element will be a planning tool that allows the scientists to schedule these userdefined observing programs to be run at a specified time and pointed at a specified target on thesun. The output of the planning tool will be instructions that can be transmitted to the instrument via
the already existing operation tools and facilities at ISAS/JAXA (section 11.3) and executed on
board.
After the observing program is run and the data is downloaded to the ground (section 11.3),
additional software routines will assemble the telemetry packets into data files that can be read
81/100using widely available computer languages. Software for “prepping” the raw data files to remove
known instrumental effects, apply the calibration, and create a higher-level data product will also be
written. Finally, basic analysis software for displaying the data and performing simple data analysis,e.g. Gaussian fits to the line profiles, will also be written and distributed to the user community.
15.7. Launch Vehicle [13-3-1]
打ち上げ機
The EUVST instrument is mounted on the standard bus. Three different spacecraft layouts wereTable 15.6: Mass budget table of the entire mission
Component

Code#

Mission section
Mission payload
EUVST

EUVST

1

X-band telemetry system

X-COM panel
Telemetry & command unit
Bus section
Management (SMS)
Satellite management unit
Data recorder
Telemetry/command IF module
Spacewire router
Mission control unit
Radio Communication (RF)S-band antenna
S-band diplexa
S-band switch
S-band hybrid
S-band transponder
Solar Array Paddle (SAP)

XMOD
XPA
XFIL
XSW
XANT

1
1
1
1
2

TCU

1

Mass (kg)Subtotal

156.8

2.4
2.5
5.2
0.1
0.3
3.0
3.5

SMU
DR
TCIM
SWR
MCU

1
1
1
2
1

2.0
1.9
2.1
3.4
1.9

SANT
SDIP
SSW
SHYB
STRP

3
2
1
1
2
2

0.5
0.8
0.1
0.1
6.7
32.4

156.8

10.5

3.0
3.5
11.3

8.2

32.4

Power Supply (EPS)Power control unit
PCU
Array power regulator
APR
SAP blocking diode
SBD
Battery (50Ah)
BAT
Attitude and orbit control system (AOCS)
Attitude orbit control computer AOCP
Reaction wheel assembly
RWA
Magnetic torquer
MTQStar tracker
STT
Inertia reference unit
IRU
Coarse sun sensor
CSAS
Ultra fine sun sensor
UFSS
Geo-magnetic sensor
GAS
AOCS interface moduleAC***
Thruster system (RCS)
RCS
Thermal control (TCS)
Heater control electronics
HCE
Thermal control materials
Structure
Instrumentation
EINT
MINT
Thruster fuelBUS dry [kg]MISSON [kg]Wet [kg]
TOTAL [kg]Figure 15.11: Spacecraft layout in the Epsilon rocket fairing
82/100

1
1
2
1

5.5
4.0
0.8
25.7

2
4
3
1
1
2
1
1
6~7
1 set

4.0
34.4
2.1
3.3
9.3
0.01
3.1

1
---

5.6
4.5

8.4
15.4

22.0
11.1

36.0

64.6

15.4
10.1

50.1
33.1

29.5
261.2
173.8
29.5
464.5

investigated with trade-off studies by a spacecraft system company as described in section 14.2(4).
Figure 15.11 shows the spacecraft layout requiring minimal modification to the bus module. The
EUVST instrument is mounted to the top panel of the bus module. The Ultra Fine Sun Sensor(UFSS) is mounted to the EUVST. The X-band module, an addition to the standard bus, is attached
to a small electric panel on the top panel of the bus module. Figure 15.11 shows the fairing
envelope, which has sufficient (20mm) clearance at the top edge of the EUVST instrument.
Table 15.6 is the mass budget table of the entire proposed mission. The mass of EUVST comesfrom section 15.8. The best estimated total mass of the mission (464.5 kg) is well within the Epsilon
capability (590kg) given in section 4 and has ample margin (24%).
Table 15.7: Instrument module design characteristics and mass breakdown estimates.
Unit/Assembly

Telescope unit
Structure
Front door
Deflector plate

Primary mirror assembly

Guide telescope
Heat rejection assembly

Description/baseline value

Precision manufactured CFRP structure, high stiffness tube type
Single open door to protect the telescope for contamination control
A plate arrangement providing an electric field to protect the mirror fromambient charged particles. A static voltage is applied across the plates
such that solar wind particles will be deflected from their path towards themirror.
Mirroroff-axis paraboloid, 28 cm diameter, 280 cm focal length,
overall < l/60 rms figure, < 5 Å rms micro-roughness,broadband VUV coating, low-expansion substrate (Zerodur)
Pointing & Fast raster/fine pointing mechanism: range 400”, imagescan
motion in pitch and yaw, <0.09” (3σ), >10 Hz response
Focus
Focus adjustment along telescope chief ray to make best
solar image on the slit.
Section 12.2.2, 0.02” (rms) accuracy at 10 Hz, range >2000”
Absorb solar heat outside FOV in front of slit. The absorber is linked to
the radiator where the heat is dumped toward the space.
Control of the mechanisms in the mirror assembly and telescope heaters.

Telescope E-box
Cable, thermal H/W, misc.
Spectrograph unit --- magnified Rowland circle design
Structure
Combined to the telescope structure which works as the optical bench forcomponents. Covered by CFRP honeycomb panel structure.
Slit assembly
5 slits on a slit plate, 0.184”, 0.368”, 0.74”, 1.47”, 2.94”, selected bylinear translation or rotating mechanism.
Slit jaws reflective to feed the light to the slit imaging assembly
Grating assemblyHouses the SW and LW gratings, both of which are Toroidal Variable LineSpace (TVLS) gratings. Focus adjustment along telescope chief ray tomake the best focused spectra on the detectors’ surface.
VUV IAPS assembly
Three 3840 × 2048, 20 µm pixel LW intensified APS detectors
bare and CsI or KBr coated MCP
IAPS electronics box
Controls and reads the IAPS detectors, provides MCP HV
EUV CCD assembly
SW CCD with radiative cooling, 2 butted 2048 × 2048,
13.5 µm pixel with shutter mechanism, thin metal foil filter
CCD electronics box
Controls and reads the CCD detectors
Slit imaging assembly
Relay optics with >200” × 300” FOV, ~0.16”/pixel, CCD or APS detector,
with shutter and filter wheel
Cable, thermal H/W, misc.
Spectrograph E-box
Telemetry and command handling, data processing, camera operations
Digital and power interfaces to the S/C
Total mass (including 15% reserves)

83/100

Mass
(kg)
87.1
43.4
3.0
1.0

22.0

3.5
1.0
3.2
10.0
69.7
31.2
1.5
1.0
7.5
4.5
1.5
2.0
4.0
8.0
8.5

156.8

15.8. Mission instrument architecture [11-3-1]
ミッションアーキテクチャ
Table 15.7 is the list of unit and assembly for the EUVST instrument and their key characteristics. A
preliminary mass breakdown (15% margin included) is also provided in Table 15.16.
An estimate of the average dissipated power for EUVST is 161 W (including 36 W for operational
heaters), which includes a contingency of 20%. In the power budget for the entire mission (623 W
in total), this value is still acceptable from viewpoint of battery balance of charge and discharge andthe upper limit capacity of the power distribution by Power Control Unit (PCU).
15.9. Ground and flight verifications [15-3-1]
地上・軌道上検証活動
Figure 15.12 shows a draft chart describing the overall flows of the flight model development. The
verification of performance, with respect to performance requirements, and the calibration of the
EUVST instrument divide into three phases.
Flight model development flow chart(not including EM/MTM etc)

Mission Instrument level
EUVST

Components

Structure (J)delivery

Instrument
AIV@Europe

Mirror assy. (J)System level

EUVST delivery

Mission – S/C Bus integration

System level test

X-COM module

S/C BusFigure 15.12 The development flow chart
15.9.1. EUVST subassembly

Prior to the instrument level AIT (assembly, integration and test), all the components will beverified by the responsible supplier. Especially, the key elements of the instrument will becharacterized at subsystem level. The key subsystems that undergo characterization are:• Mirror: optical performance and efficiency,• Gratings: the diffraction efficiency and performance at all wavelengths,• Detectors: the detection efficiencies of the cameras.
• Image motion compensation and raster system: servo control loop performance with guidetelescope
The performance of the active primary mirror assembly will be characterized at a visible
wavelength, e.g., 633 nm, using an interferometer. The resulting data will be used to verify and84/100track the expected instrument scientific performance using a model of the end-to-end performanceof the instrument. Each subsystem will be characterized by the responsible supplier.
15.9.2. EUVST assembly, integration and verifications

The EUVST instrument AIT flow is given in Figure 15.13. After subassembly verification iscomplete, the EUVST will be built up into a single structure. The first portion of the testing will be
conducted with a primary mirror simulator. The mirror simulator will allow full test of the slit jaw
imaging assembly and characterization of the spectrograph performance including cameras andelectronics. For vacuum focus and alignment, the instrument aperture will be fully illuminated using
an EUV collimator similar to Hinode/EIS. The light source is a commercially available hollow
cathode lamp. Various gas and cathode/anode combinations will generate a variety of emission lines
for image quality and throughput testing of the EUVST instrument. After the vacuum
focus/alignment of the spectrograph and slit jaw camera are complete, the primary mirror simulator
will be replaced with the primary mirror assembly. The telescope electronics box, door, guidetelescope, deflector plates and heat rejection system will be installed to complete the EUVST
instrument. The EUVST instrument performance will be verified end to end in both air and vacuum
using techniques developed for IRIS, Hinode/EIS and SO/SPICE. As shown below, the EUVST will
undergo a full suite of environmental testing at the instrument level prior to installation onto thespacecraft.

Figure 15.13: EUVST Assembly, Integration and Test flow. EUVST will be comprehensively tested
prior to integration onto the spacecraft. PM= Primary mirror
15.9.3. System Level Verification

After integrating and verifying the mission module, which includes EUVST on a support structureand X-band system on the X-COM module, the mission module is integrated with the spacecraftbus (Figure 15.13). The system level test consists of standard tests, such as electrical and functional,
vibration, thermal vacuum, and alignment. Moreover, as a special verification, micro-vibration test
85/100will be carried out during the system level test for verifying that the micro-vibration level is
acceptable to EUVST.
15.10. Acquisition surveillance: make or buy [16-3-2]
調達
JAXA is responsible for procuring the spacecraft bus and rocket vehicle and for managing theoverall development program. The EUVST instrument will be developed as an international
collaboration among Japan, US, and Europe. The telescope unit and related hardware will be
developed in Japan, whereas components installed in the spectrograph unit will be provided by aninternational consortium with institutional responsibilities given in Table 15.8. The consortium willoversee systems engineering, AIT and calibration for the spectrograph and EUVST complete
instrument level testing. The project will utilize ground support equipment for commanding andtelemetry handling at ISAS.
Table 15.8 Institute candidates of EUVST hardware contributions
1

2

3

4

5

Item
Telescope unit
Overall management
Unit structure, thermal
Aperture door assembly
Deflector
Primary mirror assembly
Mirror coating
Heat rejection assembly
Telescope electronics box
Guide telescope assembly
Telescope unit AITSpectrograph unit
Overall management
Unit structure

Unit thermal
Slit assembly
Grating assembly
VUV IAPS assembly
HVUs
IAPS electronics box
EUV CCD assembly
CCD electronics box
Spectrograph electronics box
Slit imaging assemblySpectrograph unit AITAssembled instrument
Final instrument AIT and calibration
Operations and analysis softwareHigh speed mission data downlink
Large mass memory (data recorder)
X-band telemetry sub-system
UFSS

Lead provider
JAXA
JAXA
JAXA
RAL
JAXA
MPS
JAXA
JAXA
LMSAL
JAXA

NRL
JAXA (integrated with telescope
unit)
NRL
INAF
IAS
NRL
MPS
NRL
MSSL
MSSL
LMSAL/NRLLMSAL/NRL
NRLAlternate

NRL

RAL, MPS, CSLMPSMPS (for focus mechanisms)
MSSL
MSSL
RAL, NRL, CSL
RAL, NRL, CSLMSSLNRL, MPS, CSL, JAXA for integration(may divide AIV/Calibration tasks)
EUVST consortium
Merged to Spectrographelectronics box
ESA
JAXA

15.11. WBS [19-3-1]
WBS

Work breakdown structure (WBS) is given in Table 15.9.
86/100ESA
JAXA

3000 Bus system development

Table 15.9: Work Breakdown structure2000 System design analysis

SOLAR-C Work Breakdown Structure (WBS)
Last update: 28 Dec, 2017
Prepared by SOLAR-C WG
Note: Composed at proposal submission

1000 Project management
1100 Scope management
1200 Schedule management
1300 Funding management
1400 Personnel management
1500 Risk management
1600 Reliability & quality management
1700 Configuration
1800 Communication management
1810 Level organization
1820 Project team1830 JAXA relevant departments
1840 Science commnunities
1850 Vendors
1860 Overseas Cooperation Organizations
1900 Correnpondence to phase-shift/engineering examinations
1A00 Document management
1B00 Contract & procurement management
1C00 Asset & financial management
1D00 Public relations
1E00 Correspondence to legal procedures& provisions
1E10 Radio frequency adjustment
1E20 Import & export procedures
1E30 Other legal procedures &internal provisions
1F00 Activity records

MMOD= micro-meteoroid and orbital debr5000 Bus part AIT7000 System AIT

9000 Operation

2100 Requirement analysis
3100 Bus development management
5300 Bus part development test
7100 System development test
9100 Critical phase
2200 Interface design requirement
3200 TT & C management
5310 I/F confirmation test
7110 MTM/TTM tests
9200 Initial operation
7111 Disturbance tranmission
2210 Rocket I/F
3300 Data processing system
5311 UFSS/AOCS EM fit check
9300 Normal observation operation
characteristics test
2220 Tracking control I/F
3400 Electric power system
5311 X-band/System EM fit check
9400 Post observation operation
2230 Operation & data useI/F
3500 Solar battery paddle system
5200 Bus part FM test
7200 System FM test
9500 Deorbit operation
2240 Bus system/subsystem I/F
3600 Attitude/Orbit control system
5210 Bus part subsystem FM test
7210 Integration
3610 Attitude control system component
2240 Mission/System I/F
5220 I/F confirmation test
7220 System integration test
A000 Science centerprocurement
2300 System analysis
5221 X-band/system PFM fit check7221 Circuit check5230 Attitude/Orbit control system
2310 Mission analysis
3620 UFSS development
7222 Mission system function test
A100 EUVST analysis tool
performance test
2320 Structure analysis
3620 AOCS onboard S/W development
7223 System integration test I
A200 EUVST data open access
2330 Thermal analysis
3700 Propulsion system
5230 Ground station compatibility test [E/J]
7224 Disturbance response test
2340 Mass property analysis
3800 Structure system
5300 Mission module system FM test
7225 Alignment measurement
2350 FOV & plume analysis
3900 Heat control system
5400 Bus part baking
7226 Sinosoidal vibration test
2360 Attitude/Orbit control analysis
3A00 Harness system
7227 Acoustic test
2370 Pointing error (disturbance)
3B00 Mission module system
6000 Mission part AIT
7228 Thermal vacuum test
3B10 X-band communication
analysis7229 System integration test II
system (TBD)
2380 Communication line analysis
6100 EUVST development test
722A SAP extension test
2390 Electric power analysis
3B20 X-COM module development
6110 EUVST Telescope Development Test
722B PAF fit/separation check
23A0 Reliability analysis
6111 Optical design confirmation test
722C EMC/System integration test III
23B0 Electromagnetic
4000 Mission system development
6112 Thermal optical test
722D Propulsion system leak test
6113 MTM mechanical environment test
722E Operation rehearsal testcompatibility analysis
4100 Mission system development
6114 TTM thermal equilibrium test
722F Synthetic alignment confirmatioin (final)
23C0 Environmental condition
6115 Primary mirror mechanism
management
722G Mass characteristics measurement
analysis- control system EM combination
23C1 Ventilation analysis
4200 EUVST
7
300
Launch site work
6120 Image stabilization control
23C2 Radiation analysis
4210 EUVST development7310 Shipment
system EM combination test
management
23C3 Electrostatic charge analysis
7320 Final Integration
6130 Spectrograph development
23C4 MMOD protectioin analysis
4220 EUVST system
7330 Electric test on launch site
test [E/US]
23C5 Contamination analysis
4221 Interface design
7340 Propellant filling
23D0 Operation analysis
4222 System design/analysis
6200 EUVST FM test
7350 Rocket attachment
4223 Development
23E0 Orbit analysis
6210 EUVST telescope PFM test
7360 Y-0 rehearsal
6211 Optical performanceverification plan
23F0 Propellant analysis
confirmation test
2400 System design
4230 EUVST telescope8000 Ground facilities
2410 Configuration design
4231 Telescope structure/
6212 Thermal optical test
Optical system
2420 Integration design
6213 Mechanical environment test
8100 Tracking station facilities
2430 System Distribution management4232 telescope heat control system
6214 Electric performance test
8110 Domestic stations
2431 Mass distribution management4233 Function control system
6220 Integration
8120 ESA stations (Mission data download) [E]
2432 Electric power distribution
4234 Image stabilization
6230 EUVST system integration test
8130 NASA stations (Mission data download) [US]management6231 Image stabilization control
8200 Test/Operation system(C-SODA)
control systemsystem PFM combination test
2433 Command & telemetry
4235 Deflector [E] I/F
8300 EUVSToperation system
distribution
6232 Electric performance
4240 Spectrograph [E/US]
8400 Operation room
test [E/US]
2434 Contamination management4241 Interface design
2435 Alignment management4242 Component design(incl. large volume DR & X-band)
2436 Disturbance managemen
4300 Large volume DR
6233 Optical performancef test [E/US]
2440 Safety design
4400 X-band communication system
6234 Door opening test
2450 Testing & ground test
6235 Mechanical environment test [E/US]
equipment & jigs
6236 Scientific calibration [E/US]

87/10015.12. Cost evaluation [21-3-2]
コスト見積もりTable 15.10 shows a cost evaluation on the Japanese portion of the baseline mission and flightsystem architectures. US and European costtems will be proposed to eac
cy. The US
portion (NASA) of the estimated c; the European portion isfrom theNational agencies. ESA support ofrequested. These amounts a
istent withthe expected scope of both US and Eu
n missions of opportunity.
Table 15.10: Cost details

Program Management
Comprehensive System Design
Bus System Development
S/C Management
Communication (S-band system: Low speed link, Omni antenna)
Solar Battery Paddle
Electric Power System
Attitude Orbit Control
Propelling System
Thermal Control
Mechanical Structure, Electronic Harness
Mission System Development
EUVST
EUVST-TA (Telescope Unit)
Project Management
Telescope Structure Optical system
Electronics & Drive Control
Integration, Adjustment, and Testing
Procurement of Primary Mirror
Test Equipment and Facility (NAOJ)
Test Implement Cost
EUVST-SPSpectrograph Unit
Project Management
Systems engineering
Product assurance
Slit imaging assembly
Slit assembly
Grating assembly
CCD and camera assembly
SP mechanisms (shutter, Filter wheel x 2)
IAPS detector head
IAPS camera electronics
Spectrograph electronics
SoftwareGuide telescope
Thermal for SP portion
Integration and Test for SP portion
Large Capacity RecorderJapan(10^8 JPY)

US/Europe Justificat Uncertai
(M$)
ion
nty

A
A

M
M

A
A
A
A
A
A
A
A

L
L
L
L
M
L
L
L

B
B
B
B
C
C
C
D
C, D
C, D
C, D
D
D
C
D
D
A, D
D
D
D
D
D
C,D


M
M
M
M
M
M
M
H
L
L
L
L
L
M
M
L
M
L
M
M
M
L
M


E

M

E

M

A
A
A
D

M
M
M
L

RF Compliance Test

D

L

Launch Site Operation

A

L

Maintenance of Ground Facilities

D

M

Frequency Adjustment
Travel, Labor, Public Outreach etc.
OperatioC
D

M
L

Work Consignment (System Construction & Preparation of Operation) toTechnical Dept. of Space Tracking & Communication Center (JAXA)

E

M

Initial Operation

A

L

C
C
C
E
D
D

L
L
L
M
L
M

0.00
3

0.00

--

Communication System (X-band)
X-COM Structure Panel
System Test and Flight Operation
Measurement of Disturbance Transmission Characteristics (pre-Phase A)
EM Disturbance Test
Integration Test (incl. FM Disturbance Test)
Facility Operation for System Test,

Other Items

Tracking Control Operation (2-year Stationary Scientific Observation)
Scientific Operation 2 year
Tracking at Overseas Stations (Svalbard, Troll)
Data Center Nagoya Univ.
Reserved
Rocket Procurement

0.00

Project Costexcl. Rocket, Reserved

Remarksincl. System Integration Test + UFSS 1-unit
incl. Thermal Harness

Telescope Tube: MTM assumed to be refirbished to PFM
Systems for image stabilization control and telescope operation management
30cm Mirror (FM + Backup)
NAOJ Test Facility (Extension flange for optical and thermal vacuum chamber)incl. oversea shipping costy
US ($39.05M) and European ($20.6, ESA $5.64) Cost Sharing
Based on past experience
Based on past experience
Based on past experience
Similar to IRIS, provider is the same.
Similar to Hinode/EIS mechanism
Similar to Hinode/EIS mechanism
Similar to cameras fabricated for EIS, SECCHI and SDO.
Similar to mechanisms bult for SECCHI
Similar to previous detectors for METIS/SO and Herschel SR
HVPS identical to METIS, APS electronics similar to WISPR
Similar to SECCHI/STEREO electronics box
Similar to SECCHI and IRIS software, some revuse envisioned.
Based on past experience
Based on past experience
Based on past experience
US cost sharing: included in EUVST-SP unit
Appropriate sum of system I/F equipments
X-band system: European cost sharing (ESA)
Attaching panel for X-band communication system
Measurement with exsisting bus structures
Measurement with MTM structure + EUVST (EM)Consumables for Test facilities, Liquid N2, etc. + Shipping
2 stations USC, Katsuura
X-band RF test with ESA stations (European cost sharing)
ERG-like level of Operation Rehearsal/Training, Site Operation,
Shipment for Launch Site (container reused)
Attitude control, system initial operation, ground-based computers,
and data transfer system
Support by consultants
0.2×10^8 (JPY/year) x 5 (years)

Preparation of Operation (SOOH, Operation Manuals)
Initial Operation (20 days)
Tracking Control (commander), Data Archive
Stationary Scietific Observation (2-Year)assuming (Hinode-case) International Collaboration with ESA
15% for EUVST-TA, 5% for system and operations
assuming cost of 2nd Epsilon booster, incl. JAXA-safety

Project Costexcl. Rocket
Project Cost (Total)

A: Vender's estimate, B: Based on vender's estimate, C: Hinode experience, D: Based on other mission's estimate, E: Rough estimate (no reference)

15.13. Project Organization [17-4-1]
体制
The proposed mission is a Japan-led mission with substantial participations from US and European
countries to the spectrograph unit of the EUVST. Figure 15.14 shows a preliminarily conception of
the project organization in Japan, including ISAS/JAXA, NAOJ, Nagoya U. and Kyoto U. as themain body. The EUVST instrument will be provided to JAXA by a JAXA-led international
88/100hardware development consortium (Table 15.11, Refer to supplement letters of support), and thedevelopment responsibilities reflect the expertise and strengths of each partner. JAXA will be
responsible for managing the whole instrument development. Moreover, JAXA will build thespacecraft system and the EUVST telescope unit, including the active mirror assembly, and its
accompanying spectrograph structure, and contribute to the design of the EUVST spectrograph. The
overseas part of the consortium will be responsible for managing the overall development of thespectrograph and its hardware components. The candidates of hardware contributions fromparticipating institutes are listed in Table 15.8.

Figure 15.14: Project organization
Table 15.11: International EUVST hardware development consortium
Organization

Country

Representative1
2

Naval Research Laboratory
Lockheed Martin Solar and Astrophysics Laboratory

NRL
LMSAL

USA
USA

Dr. Clarence Korendyke
Dr. Ted Tarbell

3
4
5
6
7
8
9
10

Max-Planck Institut für Sonnensystemforschung
Mullard Space Science Laboratory, UCL
Rutherford Appleton Laboratory
Institut d'Astrophysique Spatiale
Instituto Nazionale di Astrofisica
Università Degli Studi Di Padova
Centre Spatial de Liège, University of Liège
Royal Observatory of Belgium

MPS
MSSL
RAL
IAS
INAF
UNIPD
CSL
ROB

Germany
UK
UK
France
Italy
Italy
Belgium
Belgium

Prof. Sami K. Solanki
Prof. Louise K. Harra
Dr. Andrzej Fludra
Dr. Frédéric Auchère
Dr. Vincenzo Andretta
Dr. Giampiero Naletto
Dr. Serge Habraken
Dr. Andrei Zhukov

89/100Supportletter
NASA
NRL
LMSAL
DLR
UKSA
CNES
ASI
BELSPO

16.

Technical heritages, technology development status and plan
技術ヘリテージと技術開発状況と開発計画
16.1. Rationale for TRL [9-4-1]
TRL 根拠The combined EUVST technical team has decades of experience developing and operating solar
spectroscopic instrumentation. EUVST engineers and scientists have participated in thedevelopment of spectrographic solar space instrumentation since the 1970s. Technically, theinstrument represents the logical next generation instrument for spectroscopic observation of thecorona. The optical designs, mechanisms, guide telescope, slit jaw imaging system, electronics,
software and detectors have extensive heritage on previous missions. EUVST will utilize a series ofbread-board, engineering models and qualification models to demonstrate critical areas of
instrument performance. The demonstrated performance of these models is sufficient to mature theoverall instrument TRL (Technology Readiness Levels). The design and test of the mirror support
Table 16.1: TRL of key technologies and model development philosophy.
Flight helitage
Name

EUVST
Telescope
Unit

Component Name

Mirror ASY

Domestic orInternational
cooporation

DesignYellow hatched : developed by international partners

TRL

Fabrication
materialCurrent at SDR
production
process etc*Heritage from SAGEM/REOSC, Need a study for light weight andthermal deformation.
BBM(*)→PM→FM/FS*Evaluate coating samples*Based on heritage from Hinode/OTA (TRL>6 for key technologies),
BBM→MTM/TTM→PFM Need a study for thermal deformation and mechanical environment.
Require front-roading study.
*Based on flight heritage (TRL>4 for key technologies) Tip-tilt & scan
BBM→MTM/TTM→PFM mechanisms have been developed under SOLAR-C R&D (2015) andflight component for Sunrise-3/SCIP. Need launch-lock study.
* Heritage on hexapod or focus mechanism based on SOLAR-C R&D
BBM→MTM/TTM→PFM(2013-2015) (TRL>4 for key technologies)

ImportNew

Achieved

Mirror coating

Int. Coop.

New

New

4

5

Mirror supports

Domestic

New

New

2(*)

4(*)

Tip-tilt & scanmechanism

Domestic

New

Achieved

3(*)

4(*)

Focus mechanism

Domestic

New

Achieved

2(*)

4(*)

Domestic

New

Achieved

4(*)

4(*)

MTM/TTM→PFM*Heritage from Hinode OTA

Domestic
Domestic
Domestic
Domestic
Domestic
Int. Coop.
Domestic

New
New
New
New
New
Achieved
New

Achieved
Achieved
Achieved
Achieved
Achieved
Achieved
Achieved

4
4(*)
4(*)
4(*)
5(*)
9
4

4
4(*)
4(*)
4(*)
5(*)
9
4

MTM/TTM→PFM(*)
MTM/TTM→PFM
MTM/TTM→PFM
MTM/TTM→PFM
EM→PFM
PM(*)→FM
MTM/TTM→PFM(*)*CFRP based, Refurbish MTM*Based on flight heritage from Hinode OTA side door
*No need to BBM (heritage from Hinode OTA/HDM)*Based on heritage from Hinode OTA*Based on Hinode CTM-E*Build a clone of the IRIS GT*CFRP based, Refurbish MTM

Int. Coop.

New

Achieved

56Int. Coop.

New

Achieved

56Int. Coop.

New

Achieved

68

PM(*)→FMInt. Coop.

New

Achieved

6

6

BBM(*)→EM(*)→PFM

Int. Coop.

Achieved

Achieved

99

BBM(*)→PM(**)→FM/FS

Int. Coop.

Achieved

Achieved

99EM→PM(*)→FM

Int. Coop.

New

Achieved

6

6

EM→PM(*)→FM

Int. Coop.

New

Grating ASSYVUV IAPS ASSY
IAPS Electronics Box
IAPS HVPSEUV CCD ASSY
CCD Electronics Box
Spectrograph Electronics Box

6(*)

Notes

Mirror

Thermal control
structure
Unit Structure
Aperture Door ASSY
Heat Rejection ASSY
Optical Bench ASSY
Telescope Electronics Box
Guide Telescope ASSY
Spectrograph Unit Structure
Unit
Slit ASSY

6(*)

Development models

MTM/TTM (also FS)→PFM(BBM+)EM→PM(*)→FM/FS *Based on heritage from SOHO/UVCS.
EM(*)→PFM

Achieved

66

PM(*)→FM

Large mass DR
X-band telemetry sub-system

Int. Coop.
New
Int. Coop.
New
ESA or Dom. Achieved

Achieved
Achieved
Achieved

6
6
9

6
6
9

PM(*)→FM
PM(*)→FMFM

UFSS

Domestic

Achieved

4

4

BBM(*)→EM→PFM

Slit Imaging ASSY

New

*EM used to demonstrate efficiency and imaging of the spectrographsplit grating design. EM used to verify environmental performance.
*Mechanical design of large format design to be qualified.
*Modify existing WISPR/SPP design to include compression chip.
Repackage for EUVST mechanical constraints.
*If need needed at earlier phases of camera development*Identical to SO/METIS HVPS.
*Die and package previously qualified for a number of programs.
*Need to modify design to accommodate compression chip andaccelerated readout.
*Repackage existing design, add mass memory and unique software toschedule multiple spectrograph cameras*Repackage existing design from IRIS.
*Expand previous MCM designs to 39Gbits.
*Existing design from various satellites*Front-roading development (BBM) is on-going with the manufacturer
differerent from Hinode UFSS.

Yellow patched: components developed by international partners.
Reference: JAXA TRL guideline (BDB-06005 Rev.A) for details of the model and TRL definition.

Bread-Board Model (BBM) – A simple model (using commercial parts and materials) for studying key elements, suchas structures and electronic circuits, and their functions at the early phase of the design.
Engineering Model (EM) – A model for evaluating the mass, dimension, electric power and so on, in addition to theresults of functions by BBM, in order to determine the electrical and structural design.
Prototype Model (PM) – This model is subjected to qualification tests for confirming the design.
Flight Model (FM) – This model actually flies. Tested to acceptance-level testing.

Proto-Flight Model (PFM) – Two cases 1) Originally specified as prototype model and used in qualification
tests. After that, this is refurbished and evaluated as flight model, specified as the flight model, 2)Specified in advance as both prototype model and flight model. This model is subjected to a proto-flighttest (PFT) that serves both as a qualification test and as an acceptance test.

90/100structure and selection of actuators for the mirror focus mechanism will be completed early in thedevelopment plan to reduce any residual risk. The development of the mirror assembly is describedin Section 16.3.1. The EUVST team has conducted significant design studies of the mirror assemblyas described in Section 16.2.1. Table 16.1 summarizes the instrument/subassembly heritage and
EUVST model development philosophy.
16.2. Technology heritage, development history and status
技術ヘリテージと技術開発履歴と状況
16.2.1. Telescope and active tip-tilt mirror

The proposed team (NAOJ and ISAS/JAXA)
has successfully developed the 50cm aperture
visible-light space telescope for Hinode/SOT.
The telescope has a diffraction limited
performance in space (Suematsu et al. 2008),
providing 0.2-0.3 arcsec spatial resolution
images and spectro-polarimetric data of the solar
surface. This development was carried out withengineers of Mitsubishi Electric. Since then, forthe original large-strategic Solar-C mission
proposed in 2015, the proposed team with
Mitsubishi Electric engineers carried out theseries of technical feasibility studies in the Figure 16.1: Integrated concept of the EUVST
development of a 1m class solar telescope which structure.
covers near infrared, visible and UV range
widely and reached a conclusion that the technologies for the 1m class telescope are feasible to havediffraction limited optical performance, although it is challenging. In 2017, the team, with engineersof Mitsubishi Electric, has performed an initial conceptual study of EUVST, mainly from structuraland thermal point of views, to define Japanese contributions in the baseline architecture and to
identify any technical difficulty. Two structural designs were considered. The first used separate
telescope and spectrograph structures to facilitate the integration and assembly procedure. The
second incorporated a single CFRP structure integrating the telescope with the spectrograph portion(Figure 16.1). Our trade-off study found that the integrated concept better accommodated theoptical layout. Sections 12–16 of the proposal are based on the results from this study.
Technologically, one of the key areas in the EUVST concept is the primary mirror assembly, which
has a primary mirror (28 cm diameter) mounted on a tip-tilt and scanning mechanism and focusing
mechanism. A schematic design of the primary mirror assembly is shown in Figure 16.2. The
mounting of the primary mirror is one of the most critical parts in the primary mirror assembly. The
primary mirror, made of light-weighted (70% removed and thus less than 3 kg) ultra-low CTE glassmaterial, is supported by three stress-free mounting mechanisms (bipods) rooted on a tilt stage ofrotational mechanism, interfaced with three pads bonded on the side of the mirror. The pad interfaceof the mounting mechanism provides a kinematic mount for the primary mirror and avoids stresses
to the mirror resulting from dimensional errors in machining or temperature changes. Our heritageis the mounting mechanisms for the 50cm primary mirror used in Hinode/SOT (Figure 16.3), which
would be the start point for designing the EUVST primary mirror.
The tilt stage underneath the mirror consists of a two-axis gimbal mechanism for the tip-tilt image
stabilization and the slit scan observations. The tilt stage is driven with four voice-coil actuators
mounted on the mirror cell. The proposed team has a rich history developing tip-tilt mechanisms.
The team developed the tip-tilt secondary mirror mechanism for the X-ray Doppler Telescope
successfully launched by an S-520-22 sounding rocket in 1998 (Kodeki et al. 1998) and the high
91/100precision tip-tilt mirror mechanism with piezo actuators for Hinode/SOT. The image stabilization
system with this tip-tilt mechanism and servo controller achieved 0.03 arcsec (3s) on orbit(Shimizu et al. 2008), greatly contributing to diffraction-limited imaging observations for more than
11 years. For the Solar-C application, the team studied a conceptual study on the tip-tilt mechanism
for both the tip-tilt image stabilization and slit scan functions, and identified the following threetechnologies that should be studied and evaluated first: 1) sensors to measure the tilt angle with sub
arcsec accuracy, 2) low noise driver for actuators, and 3) launch lock mechanism. In 2015, the teamdeveloped a bread-board tip-tilt mirror mechanism (8 cm diameter) with built-in sensors formeasuring the tilt angle and its servo controller and confirmed tilting and scanning control withaccuracy of 0.3 arcsec (3s) (Kodeki and Shimizu 2016, and Figure 16.4 of this proposal). Thisstability performance would correspond to better than 0.05 arcsec (3s) if the same built-in sensors
are used in 28 cm diameter mechanism. Currently, the same type of tip-tilt and scan mechanism will
be used as the scan mechanism for the SCIP instrument onboard the international Sunrise balloonborne telescope (Solanki et al. 2010) which will have its third flight in 2021.
Figure 16.2:
Schematic design offocus, tip-tilt andimage scanningmechanisms for the
primary mirror
assembly.

Tilt Sensors

Rotation axes

Translational
Actuator for
Re-focusing

Primary Mirror

Mirror Support
BipodsLeaf Springs

Voice-CoilActuators for TipTilt and Slit Scan

Two-axis Gimbal
Mechanism

The mirror cell is supported by a translational
actuator like a ball-screw or a hexapod drive which
is used for re-focusing. In the case of the ball-screw
drive, the mirror cell needs to be additionally
supported by leaf springs or guide rails. In 20112014, a focus mechanism assembly (FMA, Figure16.4) was developed and qualified for usage in thespace environment for EUVST. This mechanism
uses a ball screw to translate the motor motion to thelinear motion. This bread-board model was designed
to have a step resolution of 100 µm in the range of
60 mm (Shimizu et al. 2014). For adapting it to
EUVST, higher step resolution will be required for ashorter range. As one of important verification
items, this bread-board model was successfully
operated in high vacuum environment to verify over
1x105 cycles of the back-and-forth motions (Oba et
al. 2015).

92/100Base Plate

Figure 16.3: Primary mirror with mounting
mechanisms for SOT.

Figure 16.4: Left panel shows the bread-board tip-tilt mirror mechanism while right figure displays
the translational actuator
16.2.2. Gratings

As described in section 12.1.3, EUVST gratings are based on the development history applied tovarious space instruments. Since challenging requirements will be face during the development of
EUVST gratings, the team has contacted some vendors to investigate the technical feasibility. Table
16.2 is the initial optical prescription of the LW and SW gratings. Similar to SoHO/CDS, these two
grating halves are mounted in close proximity. The gratings are figured and coated to optimize theimage quality and efficiency of each passband. The highly corrected toroidal gratings with variable
line spacing give excellent image quality at the focal plane across a large wavelength range. The
grating efficiency performance has been modeled and is presented in Figure 16.5.
Table 16.2: Grating Optical Characteristics
Grating Characteristic
First Order Wavelength Coverage
Grating Size (illuminated area)

RMS figure error
RMS surface roughness
Material
Aspect ratio (full diameter to thickness)
Groove Efficiency
Blaze
Coatings

Short Wavelength TVLS
17 nm – 21.5 nm
Half of 50mm diameter with 2
mm edge margin< λ/64<0.5nmFused Silica
5:1>40% average
Laminar
Graded Mo/Si multilayer

Surface geometory: radii of curvature
R1: dispersion direction
Nominal Angle of Incidence (degrees)
Central Ruling Density
Order of Diffraction

R1= 587.8836 mm
R2= 545.74 mm-3.54200 lines/mm
First

93/100Long Wavelength TVLS
69-128nm
Half of 50mm diameter with
2mm edge margin< λ/64<0.5nmFused Silica
5:1>40% average
Triangular blaze
Tb-W-Si multilayer with B4C
capping layer
R1= 475.8877 mm
R2= 503.13 mm-2.1
2000 lines/mm
first and second

Figure 16.5: Left panel shows the modeled multi-order performance of the SW grating (including
the laminar groove efficiency and the SW coating) while right panel displays the LW grating(including the triangular blaze and the multilayer coating). An opaque layer of B4C was alsomodeled for the LW grating.
16.2.3. Coating

Coatings for the EUV and FUV spectral ranges have highly mature designs and extensive flight
heritage. EUV multilayer coatings have been developed, implemented, qualified and flown for
instruments such as SoHO/EIT, TRACE, SDO/AIA, STEREO/SECCHI, Hinode/EIS, the EUNIS
and Hi-C rocket flights. Recently, the qualification phase has been successfully completed for
SO/EUI and SO/SPICE (B4C thin layer on quartz substrate). The coatings needed by EUVST (seeTable 16.3) are based on what was already qualified and flown. The coating design will be modified
to incorporate the broad bandpass required for the EUV and the B4C thin cap layer for enhanced
performance for the main mirror coating.
Table 16.3: The coatings for EUVST
Opticalcomponent
Main Mirror

SW Grating

LW Grating

Required coating

Heritage

Development status

Broad-band multilayer to cover
the 17 to 21.5 nm range with
B4C thin cap layer to provide
reflectivity in the 46 to 127 nm
range.
Broad-band multilayer to cover
the 17 to 21.5 nm range.

SoHO/EITTRACE
STEREO/SECCHI
Hinode/EIS
SDO/AIAEUNIS

B4C thin cap layer to provide
reflectivity in the 46 to 127 nm
range.

SO/SPICE

Feasibility study with sample production
and reflectivity measurements for thebaseline Mo/Si multilayer with thin B4C
cap layer carried out by consortium
members.
Feasibility study with sample production
and reflectivity measurements for thebaseline Mo/Si multilayer carried out by
consortium members.
Mature design.

A feasibility study of a suitable multilayer coating (Mo/Si with B4C barrier layers) with and without
B4C thin cap layer (for the main mirror and for the SW grating, respectively) was successfullycompleted and showed:
1) the coating shall be resistant to temperatures up 120°C, ideally up to 200°C.
2) micro-roughness of the surface should not be worsened more than 0.2 nm rms by the coating,
The study included coating different samples with various thickness of B4C cap layer (7, 11 and15nm), thermal cycling up to 100 h up to 200 °C, atomic force microscopy measurements of microroughness and VUV reflectivity measurements. The EUVST team is also studying an Al/Mo/SiC
coating with enhanced reflectivity (25%) at shorter wavelengths and a significantly lower solar
absorption (Figure 16.6). Similar coatings were used on EUI on Solar Orbiter. This coating will
increase reflectivity ~25% over the entire wavelength range.
94/100Figure 16.6: The reflectivity in UV and visible spectrum of an Al/Mo/SiC coating recently
designed.
16.2.4. Ultra Fine Sun Sensor

The Ultra Fine Sun Sensor (UFSS) has been working withhigh precision on the Hinode satellite for over 11 years on
orbit. The sensing accuracy was vastly improved over similarsensors on Yohkoh and Akari. However, the manufacturer nolonger provides the UFSS. In 2013, we identified an alternate
manufacturer and started to develop a second generation of
UFSS with the participation of the chief engineer (Dr. Tsuno,
currently at RIKEN) responsible for the original UFSS. The
new development basically follows the design of the first
generation UFSS but a couple of upgrades have been made.
Specifically, the analog circuit designs were changed todigital. In 2014-2017, we carried out various developmenttasks, including the following: the performance of chosen 1dimensional CCD sensors with radiation exposure were
tested; each term in the error budget for achieving accuracy
was evaluated; a bread-board model circuit was designed andevaluated with a bread-board model (Figure 16.7 top); thereticle glass (Figure 16.7 bottom) was evaluated; and the
Figure 16.7: (top) Electrical circuit
structural design and thermal deformation analysis wasand (bottom) reticle developed forcompleted. The TRL has almost reached 4 in 2017. Currently,
UFSS performance evaluation.
a structure is under fabrication for experimentally validating
the expected thermal deformation. This structure will be also used to evaluate sensing performanceby integrating the bread-board electronics model to reticle-CCD in a flight like structure.
16.2.5. Micro-vibration evaluation

Using the heritage accumulated in the development of the Hinode/SOT, we developed a successful
strategy for controlling the micro-vibration for EUVST. The following approach will be adopted ineach development phase:

1) Design phase
Identify the sources of the mechanical disturbance and evaluate the level of disturbance thatthey produce using the data of identical devices in hand. Identify the optical components that are
95/100critical for the image stability and evaluate their sensitivity to the image shift. Calculate thetransfer function of mechanical disturbance using the preliminary structural models of the bus
module and EUVST, and predict the image motions on the image plane to identify the criticalareas. Study the methods to mitigate the unacceptable vibration of optical components if it isobserved in the measurements in later phase.
2) Component fabrication and procurement phase
Control the disturbance levels of individual disturbance sources during their fabrication withvendors of each component.
3) Development and testing phase
Using the MTM (or dummy structures) of the bus module and EUVST, measure the transfer
function of mechanical disturbance from the disturbance sources to the optical components. In
the measurement, a shaker that produces translational and torque forces is attached at thelocation of the disturbance sources and accelerometers are attached on the critical opticalcomponents to simulate the disturbance sources and to evaluate the image motion, respectively.
Using the obtained data, improve the structural model to better predict the image jitter, and, if
unacceptable level of image jitter is anticipated, the method for mitigating it will be applied.
4) Flight model phase
Measure the end-to-end performance of image stability using the integrated flight model.
So far, we have identified the primary sources of mechanical disturbance as the 4 reaction wheels(RW) and one IRU unit containing 3 gyros, and the critical optical components for the image
stability as the primary mirror, grating and slit of EUVST. Transfer functions from the 5 disturbancesources to the 3 optical components were evaluated using the preliminary structural model and the
disturbance data of Honeywell’s RWs and Type-II IRU of Mitsubishi Precession Corporation. In
each combination of the disturbance source and the optical component, the transfer function was
calculated for plausible injected disturbances of translational forces in x, y, and z and of torqueforce in qx, qy, and qz, and for the resultant motions of the optical component in lateraldisplacement in x, y, and z and rotations in qx, qy, and qz. Thus, 6 x 6 x 4 x 3 = 432 transfer
functions were calculated in total in the frequency domain of 1-300Hz. Our preliminary study infers
that the disturbance produced by the IRU unit is about one order of magnitude smaller than therequirement, but the RWs could produce unacceptable level of image jitter in the frequency range of
50-110Hz. Note that a conservative magnification factor of Q=400 is assumed in this study for thesafety, while, even though the disturbance level of the IRU is small, the IRU unit may produce asignificant image jitter once a resonance in the real hardware occurs at 155Hz. The accuracy of theprediction needs to be improved using the data obtained from the measurements with the dummy
structures and MTM.
As for the methods for mitigating the potential risk on micro-vibration, we will use the following
approach if unacceptable disturbances are observed during development phase.
1) Insert a vibration isolation damper in the path between the disturbance sources and EUVST.
The dumper will be designed to reduce the transfer function by a factor 10 ~ 100 at thefrequency range of f >50Hz. A possible location of the dumper will be the mechanical interface
between the bus module and the EUVST telescope. Note that isolating the 4 RWs in the bus
module is not easy because it requires a significant modification to the design of the standard
bus module.
2) Reduce the disturbance level of RWs and IRU
This needs to be worked out with the vendors of both components. In addition to thespecification control by documents, accurate measurement of the disturbance levels of the real
hardware, fine tuning of the mass balance of the wheels and screening of the devices will be
required.
3) Limit the operational range of the spin rate of RWs
96/100Since the amplitude of the dominant disturbance of RWs increases as the square of the spin rate,
limiting the spin rate in the low rotational range (e.g. < 3000rpm) will contribute to reduce themicro-vibration. Resonance frequencies of optical components obtained from the measurement
will be avoided. Possible drifts of the spin rate of RWs in orbit need to be studied.
4) Revisit the requirement of the image stability and/or the allocation of allowable disturbance toeach component.
16.2.6. Contamination control in UV

A systems engineering approach on the contamination control was performed in the Hinode (SolarB) development program. It contains the material selections, the sample outgassing measurementswith the Thermoelectric Quartz Crystal Microbalance (TQCM), the contamination mathematicalmodeling activity, personnel training, parts cleaning, the bake-out for reducing outgas, outgassing
measurements of the flight models with TQCM, contamination monitoring of flight models withwitness plates in the subsystem and system integration and test activity. The throughput degradationof the optical telescope on Hinode has happened with a slightly higher rate than the predicted,
especially at the short wavelength range of ~400 nm due to the molecular contamination. A further
careful control needs to be applied in the EUVST program that requires a higher-level control thanthat in Hinode. More difficult contamination control has been applied to the SUMER instrument on
SoHO and EUI, SPICE, and Metis on Solar Orbiter (SO), those of which function in the UV
wavelength range that EUVST shares. While the metal structure was used in SUMER for an easier
treating of the molecular contamination control, the structures of the SO science instruments arebuilt with CFRP materials for light-weight payload requirement, as EUVST does. The heritage of
the SO contamination control is to be fully used. The primary difference from the Hinode’s
treatment is the high-temperature bake-out of the payload parts under a low vacuum condition of afew Torr during a low-rate GN2 purging in an oven for effective cleaning before a high vacuum
bake-out in 10-6 Torr.
Pre-Phase A1b/A2
TelescopePhase AFY2018Phase BFY2019FY2020(Mission) SDR

I/F definition

Mirror assembly
Assembly feasibility study
Assembly BBM design
Evaluation in computers

item

Mirror supports
Thermal deformation

between US/Europe partners and telescope
between EUVST and system

MTM design

Assembly BBM evaluation
Assy. Manufacture

Tests

Mirror specification for MTM

Fabricate MTM mirror (lead: 12mo)

Focus mechanism
Application
Tip-tilt mech.

Application
Launch-lock mechanism

Microvibration
Prediction with model fidelity update

UFSS
BBM evaluation

Design/evaluation
MeasureQ-value of the Bus(with dummy EUVST structure)
Define AOCS I/F

Figure 16.8: Technology development schedule
97/100Evaluate transfer
function(with EUVST MTM)
EM design16.3. Technology development plans
技術開発計画
Table 16.1 shows a draft of the model development strategy for assemblies and units in the mission
payloads. Figure 16.8 describes the overall development for Japanese portions of the technology,i.e., telescope unit and its mirror assembly, micro-vibration control, and UFSS. Details of eachdevelopment plan is explained in the subsequent subsections. In Phase A, we will take efforts fordefining the interfaces. These efforts include the studies and negotiations to define the interface
between the telescope and US/European partners as well as between the EUVST instrument and thespacecraft system. For some critical technologies in Figure 16.7, we will make R&D efforts in Prephase A1b/A2 phase.
16.3.1. Strategy for developing the mirror assembly

The primary mirror assembly will be developed as a single assembly based on three key
technologies, i.e., mirror supports developed for Hinode, tip-tilt mirror mechanism and focus
mechanism which have been studied for a Solar-C telescope.
Pre-phase A1b/A2 is the time to take an assembly design for bread-board model to evaluate thethermal deformation performance and its feasibility evaluation based on model calculations. The
thermal deformation of the mirror surface is required to be minimized at the mirror temperature ashigh as +80°C. The baseline design of the mirror supports and mirror pad that holds the glass mirror
were developed for the 50cm primary mirror of Hinode/SOT. The mirror supports use bearings torelease the moment properly and give less deformation to the surface figure when the mirror ismounted on the assembly as well as when the temperature is changed significantly. The mirror
supports must survive the launch environment and be light weighted to minimize the weight of thesteering portion. The mirror pad uses an Epoxy adhesive to contact the mirror, but the outgas
properties are one of key portions in design. In case that the adhesive does not meet the outgas
requirement, we will consider a backup non-adhesive support method, which has been used as themirror supports for the Hinode/SOT tip-tilt mirror mechanism. The multilayer coating may also
have internal stress and deform the surface figure. The measurement of the coating stress will be
used to evaluate the deformation caused by the coating. After evaluating the thermal deformation of
the mirror supports with computer software, we will manufacture a bread-board model in the later
period of pre-phase A2 and phase A, and experimentally verify the results from the model
calculations. The bread-board model consists of three mirror supports with a dummy mirror, which
has sensors for measuring the stress. The important milestone in the development of the mirror
assembly is the start of the MTM mirror fabrication by the mirror company. The lead time of amirror is long (between 12-18 months, depending on the specification). By the beginning of FY
2020, the specification of the mirror should be defined and agreed to keep the schedule proposed in
section 4.
The structural design concept for the tip-tilt and scan mechanism as well as the focus mechanism inthe mirror assembly will be studied in pre-phase A1b/A2. The study includes a couple of technicaltrade studies to choose the best solution from the existing technologies, such as pivot and flexure
for steering. A trade-off study for actuators and sensors was carried out in 2015 R&D development
for the Solar-C 1m telescope, but this study should be revisited for the EUVST to define thebaseline design parameters. The focus mechanism to be designed for the mirror assembly uses
results from R&D in 2011-2014 as the baseline, but it is important to confirm whether thedeveloped ball-screw mechanism is applicable to the mirror assembly in pre-phase A1b/A2. Thesedesk studies will give inputs to the structural design included in a bread-board model of the mirror
assembly in phase A. A launch lock mechanism will be needed to hold the steering portion during
the launch. The lock mechanism will be investigated in pre-phase A1b/A2 with help of spacecraft
98/100systems and JAXA mechanical experts. Our plan is to adapt a non-shock type actuator recently
developed to the launch lock in the mirror assembly.
16.3.2. Strategy for developing the UFSSThe R&D efforts with bread-board model have been in progress as described in section 16.2.4. The
remaining major items to be evaluated are: 1) experimental evaluation of the thermal deformationby using the structure fabricated in FY2017, and 2) experimental evaluation of the performance(random noise and bias noise) by combining the bread-board electronics with the reticle-CCD built
in the structure. These works will be carried out in Pre-phase A1b/A2. The interface with theattitude control system will be discussed in Phase A before going to the EM design.
16.3.3. Strategy for performing the micro-vibration evaluation

Micro-vibration evaluation will be carried out in the following steps:
1) Pre-phase A1b/A2
Our preliminary study summarized in section 16.3.2 was carried out by NEC engineers, assuming
the EUVST structure which is in form of separated telescope unit and spectrograph unit. As shownin Figure 16.1, an EUVST study has suggested that the integrated structure is preferable. According
to further updates on optical and mechanical designs in Pre-phase A1b/A2, the mechanical model ofthe EUVST will be updated to improve the model fidelity and thus the prediction of the microvibration. For mitigating risks of the micro-vibration and of the procedure of the micro-vibration
testing, it is important to define details of verification plans in this phase.
2) Phase A
Using the existing structure of the ISAS small bus module with a dummy EUVST structure, weplan to perform an evaluation test. This test will measure Q-values of the transfer functions from thelocations of RWs and IRU to the location of optical components in EUVST. Especially, this will be
the first opportunity to know transfer functions inside the bus module. A small shaker is used as a
disturbance source and the level of micro-vibration will be measured with accelerometers attachedto the EUVST dummy structure. The results from the measurements are incorporated in themathematical model to improve the prediction. The supporting structures of EUVST will be
designed, including the necessity of a vibration isolation damper, based on this prediction.
3) Phase B
Using the existing structure (or MTM if needed) of the ISAS small bus module with a MTM modelof the EUVST, we will measure transfer functions of the mechanical disturbance from the locationsof RWs and IRU to the optical components in EUVST. A shaker is used as a disturbance source,with accelerometers attached on the critical optical components. The results from this test areincorporated in the mathematical model and we will identify whether any modification is needed forthe final design of the EUVST supporting structures to avoid any type of resonance between the
disturbance sources and the optical components.
4) Phase D
In the final spacecraft test, we plan an end-to-end testing of micro-vibration using the integrated
flight model of the spacecraft and confirm the amplitude of the image jitter to be within the required
level. Test procedures to measure the amplitude of the image jitter will be studied in Pre-phase
A1b/A2. Based on the results from this final test, we will finalize the operational range the spin rateof RWs.
16.3.4. International contribution items

1) Gratings
Two vendors (Horiba JY and Zeiss) have indicated that they can provide holographic TVLS
gratings for EUVST. We plan to order prototypes during the formulation phase of the mission inorder to validate the detailed grating specifications (groove profile, depth, density, etc.). The
99/100gratings will be thoroughly tested in the EUV prior to the start of Phase C to verify groove
efficiency and imaging performance. As a backup, the EUVST team has considered the
procurement of a spherical VLS grating and a mechanically ruled VLS grating from Bach
Associates. Both backup options are viable and would satisfy the science threshold requirements.
The backup designs will accommodate the existing mechanical interfaces. The backup options will
only be exercised if the prototype gratings prove to have unacceptable performance.
The grating coatings multilayer technology has flown on numerous imagers (SoHO/EIT,STEREO/EUVI, TRACE, SDO/AIA, etc.) and spectrometers (SERTS, SoHO/CDS, Hinode/EIS,VERIS). The baseline design of the SW coating is a very robust Mo/Si multilayer design. Alternate,
more efficient designs based on other combinations of materials will be studied during phase A.
Structures giving more than 50% reflectivity around 19 nm include Al/Mo/B4C and Al/Mo/SiC, as
used on Solar Orbiter/EUI (Delmotte et al. 2013).
2) Coating for the primary mirror

The feasibility study summarized in section 16.2.3 shows the readiness and adherence torequirements of the baseline Mo/Si multilayer based the design. On the other hand, simulations run
at the Institute of Optics (based on the coatings developed for Solar Orbiter/EUI) show beneficial
effects from the addition of aluminum layers to the multilayer formula in terms of increased
reflectivity at visible and IR wavelengths (Auchere, private communication) that translates in alower equilibrium temperature of the main mirror. During phase A, further studies will be conductedin order to fine-tune the multilayer response in terms of reflectivity in the selected science bandpasses as well as at visible and IR wavelengths to minimize the heating of the mirror. The
optimization will include sample production and reflectivity and emissivity measurements. The
realization of the flight coating will follow a BB (samples), PM and FM/FS approach on samples
and optics provided by the mirror assembly responsible.

100/100