maruma/spica/old/MDD_090828.doc · Web viewPLM Payload Module PM Prototype Model POM (SAFARI) Pick-OFF Mirror [Q] QT Qualification Test [R] RCS Reaction Control Subsystem ROIC Readout

　　　　　 SPICA MDD　Reference: JAXA_SPICA_SYS_MDD_001

Date: 5/6/2023　

Issue: 1.0 　 Page:　i

SPICAMission　Definition　Document

Issue 1.0 (Aug.27, 2009)


Date: 5/6/2023　

Issue: 1.0 　 Page:　ii

Document　Change　Record

Issue/Rev. Date Version　 Page　Affected

1.0 27 Aug, 2009 Initial Issue for Down

Selection of ESA Cosmic

Vision

New Document


Date: 5/6/2023　

Issue: 1.0 　 Page:　iii

Table　of　ContentsList　of　Acronyms.................................................................................................................................1

1.　Introduction....................................................................................................................................7

1.1　Purpose　of　the　Document.........................................................................................................7

1.2.　Overview　of　SPICA.................................................................................................................7

2.　Project　Status..................................................................................................................................9

2.1.　Status　in　Japan.........................................................................................................................9

2.2.　Status　in　Europe.....................................................................................................................10

3.　Summary　of　SPICA　Science　Objectives......................................................................................12

3.1　Overview..................................................................................................................................12

3.2　Resolution　of　Birth　and　Evolution　of　Galaxies.....................................................................12

3.3　Revealing　the　Transmigration　of　Dust　in　the　Universe........................................................13

3.4　Thorough　Understanding　of　Planetary　System　Formation.................................................14

4.　Mission　Requirements..................................................................................................................15

4.1　Focal　Plane　Instrument　Specifications..................................................................................15

4.　2　Pointing　Control　Requirement.............................................................................................17

5.　Mission　Profile..............................................................................................................................19

6.　Spacecraft　Design.........................................................................................................................25

6.1　Overview..................................................................................................................................25

6.1.1　Layout...............................................................................................................................25

6.1.2　Satellite　Mass　and　Electric　Power..................................................................................27

6.2　Mechanical　Structure.............................................................................................................29

6.3Thermal　Control......................................................................................................................31

6.3.1　Thermal　Environment.....................................................................................................31

6.3.2　Thermal　Design................................................................................................................31

6.3.3　Thermal　Conditions　at　the　Interface　with　Payload　Module.........................................32

6.4　Electrical　Power　Subsystem...................................................................................................33

6.4.1　Overview...........................................................................................................................33


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6.4.2　Solar　Array　Paddle　(SAP)...............................................................................................34

6.4.3　Shunt　Dissipater　(SHNT)................................................................................................34

6.4.4　Power　Control　Unit　(PCU)..............................................................................................35

6.4.5　Battery　(BAT)...................................................................................................................35

6.4.6　Performance.....................................................................................................................36

6.5　Communications　Subsystem..................................................................................................37

6.6　Data　Handling.........................................................................................................................41

6.6.1　Architecture......................................................................................................................41

6.6.2　Functions..........................................................................................................................42

6.6.3　Performance.....................................................................................................................43

6.7　Attitude　and　Orbit　Control　System　(AOCS)........................................................................44

6.8　Reaction　Control　Subsystem..................................................................................................48

6.8.1　System　Description..........................................................................................................49

6.8.2　Performance.....................................................................................................................49

7.　Payload　Configuration.................................................................................................................52

7.1　Cryogenic　System...................................................................................................................52

7.1.1　Thermal　Insulation　and　Radiative　Cooling　System　(TIRCS)......................................52

7.1.2　Mechanical　Cooling　System　(MCS)................................................................................56

7.2　STA..........................................................................................................................................58

7.3　FPI............................................................................................................................................62

7.3.1　Overview...........................................................................................................................62

7.3.2　Mid-Infrared　Camera　:　MIRACLE...............................................................................65

7.3.3　Mid-IR　Spectrometer:　MIRMES　&　MIRHES..............................................................67

7.3.4　SPICA　Coronagraph　Instrument:　SCI..........................................................................69

7.3.5　Focal　Plane　Finding　Camera..........................................................................................71

7.3.6　SAFARI.............................................................................................................................72

7.3.7　BLISS................................................................................................................................73

8.　AIV/T　Plan....................................................................................................................................75

8.1　General....................................................................................................................................75

8.2　STA　in　Japan...........................................................................................................................76

8.3　FPI............................................................................................................................................78

9.　Project　Plan...................................................................................................................................80


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Issue: 1.0 　 Page:　1

List　of　Acronyms

[A]

ACC Accelerometer

ADR Adiabatic Demagnetization Refrigerator

ADS　 Attitude Determination System

ALFRP Alumina Fiber Reinforced Plastics

AGB Asymptotic Giant Branch

AGN Active Galactic Nuclei

AIV/T Assembly, Integration and Verification/Test

AOCP Attitude and Orbit Control Processor

AOCS Attitude and Orbit Control Subsystem

ASIC Application Specific Integrated Circuit

AT Acceptance Test

[B]

BAT Battery

BBM Bread Board Model

BLISS Background-Limited Infrared Submillimeter Spectrograph

BM Bus Module

[C]

C-FPC Focal Plane Camera for Coronagraph

C-TTM Tip-Tilt Mirror for Coronagraph

CCSDS Consultative Committee for Space Data Systems

CDHU　　　　　　　Central Date Handling Unit

CDM Cold Dark Matter

CFRP　 Carbon Fiber Reinforced plastics

COR Coronagraph

CPW Coplanar Waveguide

CTE Coefficient of Thermal Expansion

CTIA Capacitive Trans-Impedance Amplifier


Date: 5/6/2023　


[D]

DI Direct Injection

DPU Digital Processing Unit

DRV Drive Unit

[E]

ECSS European Cooperation for Space Standardization

EDAC Error Detection and Correction

EGSE Electrical Ground Support Equipment

EM Engineering Model

EMC Electro-Magnetic Compatibility

EMI Electro-Magnetic Interference

EOL End of Life

EPR Ethylene Propylene Rubber

EPS Electric Power Subsystem

ESA European Space Agency

ESI European SPICA Instrument

　

[F]

FDIR Fault Detection Isolation and Recovery

FDM Frequency Domain Multiplexing

FIR Far Infrared

FM Flight Model

FOV Field of View

FPA Focal Plane Array

FPC-G Focal Plane Camera for Guidance

FPC-S Focal Plane Camera for Scientific Observation

FPGA Field Programmable Gate Array

FPI Focal Plane Instrument

FPI-E Focal Plane Instrument Electronics

FPSTT (Same instrument as FPFC)


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FPU Focal Plane Unit

FSS Fine Sun Sensor

FTS Fourier Transform Spectrometer

[G]

GEMS Glass with Embedded Metals and Sulfides

[H]

HGA High-Gain Antenna

HIFI Heterodyne Instrument for the Far Infrared (Herschel Instrument)

HIRES (MIR) High Resolution Spectrometer

HK House Keeping

HRS Heat Rejection System

[I]

ICU Instrument Control Unit

IFTS Imaging Fourier Transform Spectrometer

IFU Integral Field Unit

IOB Instrument Optical Bench

ISAS Institute of Space and Astronautical Science

[J]

J-T Joule-Tompson

JAXA Japan Aerospace Exploration Agency

JWST James Webb Space Telescope

[K]

KBO Kuiper Belt Object

KID Kinetic Inductance Detector

[L]

LEKID Lumped Element Kinetic Inductance Detector

LNA Low Noise Amplifier

LWS (MIR) Long Wavelength Spectrometer


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[M]

MAG Magnetometer

MCS Mechanical Cooling System

MGSE Mechanical Ground Support Equipment

MIR Mid infrared Instrument

MIRACLE Mid-Infrared Camera w/o lens

MIRHES Mid-IR High-Resolution Echelle Spectrometer

MIRMES Mid-IR Medium-Resolution Echelle Spectrometer

MKID Microwave Kinetic Inductance Detector

MLI Multi-Layer Insulation

MOC Mission Operation Center

MRD Mission Requirement Document

MSFC Marshall Space Flight Center

[N]

NEP Noise Equivalent Power

[O]

OBS On-Board Software

OGSE Optical Ground Support Equipment

OPD Optical Path Difference

OWA Outer Working Angle

[P]

PACS Photo Detector Array Camera and Spectrometer (Herschel Instrument)

PAF Payload Attach Fitting

PC Photoconductor

PCU Power Control Unit

PFM Proto-flight Model

PIAA Phase Induced Amplitude Apodization

PLL Phase Locked Loop

PLM Payload Module


Date: 5/6/2023　


PM Prototype Model

POM (SAFARI) Pick-OFF Mirror

[Q]

QT Qualification Test

[R]

RCS Reaction Control Subsystem

ROIC Readout Integrated Circuit

RTU Remote Terminal Unit

RW Reaction Wheel

[S]

S/C　　 Spacecraft

SAFARI SPICA Far-Infrared Instrument

SAP Solar Array Paddle

SCI SPICA Coronagraph Instrument

SDR System Definition Review

SED Spectral Energy Density

SEMP SPICA System Engineering Management Plan

SEU Single Event Upset

SHUNT　　 Shunt Dissipater

SMBH Super-Massive Black Hole

SNR Super-Nova Remnant

SOAP Sate-of-Art Performance

SOC Science Operation Center

SPICA Space Infrared Telescope for Cosmology and Astrophysic

SRR System Requirement ReviewSSMM Solid State Mass memory

STA SPICA Telescope Assembly

STT Star Trackers

SVM Spacecraft Service Module


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SciRD SPICA Telescope Science Requirement Document

SiC Silicon Carbide

[T]

TAC Telescope Allocation Committee

TBA To Be Agreed

TBD To Be Decided

TDM Time Domain Multiplexing

TES Transition Edge Sensor

TIRCS Thermal Insulation and radiation Cooling System

TM Telemetry

TNO Trans-Neptunian Object

TOB Telescope Optical Bench

[U]

ULIRG Ultra luminous IR Galaxy

[W]

WE Warm Electronics

WFE Wave Front Error

[X]

XRCF X-ray Calibration Facility


Date: 5/6/2023　


1.　Introduction　

1.1　Purpose　of　the　Document

The purpose of the Mission Definition Document (MDD) is to provide an overview of the SPICA

mission and to define and control the overall interface ranging from the SPICA spacecraft to each of

the SPICA scientific instruments. MDD describes the implementation of the instrument requirements

in the design of the SPICA spacecraft and is a result of the spacecraft design activities performed by

JAXA and contractors.

The current MDD (ver.1.0) is prepared for the internal review by ESA for the down selection under

the framework of the ESA Cosmic Vision. Hence the current MDD is not a public document, and can

be distributed only to ESA reviewers. If the whole document or part of the MDD is to be used

outside ESA, the permission by JAXA is required prior to its use.

1.2.　Overview　of　SPICA

SPICA (Space Infrared Telescope for Cosmology and Astrophysics) is an astronomical mission to

reveal the evolutionary history of the universe, ranging from the birth and evolution of galaxies, the

formation and evolution of stars and planetary systems, to the chemical evolution of the universe. On

the basis of the AKARI and Herschel heritages, SPICA aims to address a variety of key issues in the

current astrophysical problems.

Fig.1.1 Artists impression of SPICA in orbit


Date: 5/6/2023　


With a 3m class cooled (nominal <6K, target <5K) telescope, SPICA enables observation with the

unprecedented spatial resolution and sensitivity. To carry the massive cooled telescope, we have

employed a Warm Launch System, in which the telescope is to be launched at ambient temperature

and cooled down in orbit.

International collaboration on the SPICA mission has been discussed extensively. Especially,

European participation to the SPICA project was approved as a candidate of ESA future missions

under the framework of the ESA Cosmic Vision 2015-2025.

The target year of the launch is 2018.


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2.　Project　Status　

2.1.　Status　in　Japan

SPICA mission concept is based upon extensive discussions over 10 years among the astronomical

community in Japan, and the SPICA project is now at the official “pre-project” status at JAXA. The

SPICA project is required to take one more major step for the final approval.

In 1999, the SPICA working group was established under the Space Science Committee of Institute

of Space and Astronautical Science (ISAS). The group members were mainly from the AKARI (the

first Japanese infrared space telescope satellite) community and the SUBARU telescope user

community and also include astronomers from various fields reflecting great interests on the project

from the communities. The working group studied the mission concept, and also made extensive

activity on the development of key technologies which are indispensable to enable the mission.

In 2005, the National Committee for Astronomy, Science Council of Japan reviewed the SPICA

mission proposal and strongly recommended it as one of the most important missions in the Japanese

space astronomy program.

In September 2007, the SPICA working group submitted a new mission proposal (2nd edition) to

ISAS. Following this, SPICA went though the Mission Definition Review (MDR) and was approved

in the procedure by the Space Science Steering Committee of ISAS in March 2008.

In May 2008, SPICA went through the Pre-Project Review, which is a management review by

JAXA directors. The review result was approved and the SPICA pre-project team was established

officially at JAXA on July 8, 2008 by the President of JAXA. This means the SPICA pre-project

team’s activity is fully approved at least during Phase-A and one of the purposes of the team is to

prepare for the next management review, the Project Approval Review, for the realization of the

project.

The SPICA schedule from Phase-A to the beginning of operation is shown in the figure 2.1. Phase-

A consists of two smaller phases, Concept Development Phase and Project Formation Phase. In the

Concept Development Phase, we will define requirements for the whole system, which will meet the

Mission Requirement Document (MRD). The goal of the Concept Development Phase is to get

approval in the System Requirements Review (SRR), which is expected to be held at the end of

FY2009. JAXA SPICA pre-project team is working with three industries in Phase-A: Sumitomo

Heavy Industry (SHI), NEC and MELCO. SHI is the only domestic industry which has advanced

space Cryo technology. Two others are the satellite system experts, each of whom is working with


Date: 5/6/2023　


JAXA individually. JAXA has contractual relationships with the two industries to obtain cooperation

in the R&D from different approaches. NEC was the contractor of AKARI (ASTRO-F) satellite.

MELCO was the contractor of HINODE (Solar-B).

In the Project Formation Phase, we will define the details of the whole system, which will meet the

systems requirements defined in SRR. The SPICA pre-project team is required to present not only

the science and technical design, but also the project and engineering plan to demonstrate the high

feasibility of the mission. The goal of the Project Formation Phase is to get approval in the System

Definition Review (SDR), which is expected to be held in early FY2011.

At the end of Phase-A, SPICA will go through the Project Approval Review, which is a

management review by JAXA directors. This review makes an important decision on whether

SPICA can proceed though phase-B to the end of mission. The review is expected to be held by mid

2011 to meet the launch target of the SPICA in 2018.

The SPICA mission is the international collaborative project of JAXA and ESA. Wining the

Cosmic Vision down selection is a necessary condition for getting approval in JAXA’s management

review. Thus the SPICA pre-project team and ESA SPICA team should work in close cooperation,

holding monthly teleconferences to share the progress of studies and situation in both of them.

2.2.　Status　in　Europe

The European SPICA Consortium (P.I.: B. Swinyard, RAL, UK) submitted a proposal to enable

European participation in SPICA to ESA in June 2007 under the framework of the ESA Cosmic

Vision 2015-2025. The proposal called on ESA to assume a partner agency role in SPICA by making

the contributions in: (1) SPICA Telescope Assembly, (2) European SPICA Ground Segment, (3)

SPICA Far-Infrared Instrument System (SAFARI) Engineering and Management, and (4) SPICA

Mission support. The proposal also assumed that SAFARI was to be developed by the European

Consortium.

The proposal was selected by ESA in October, 2007, as one of candidates for future missions.

Following this, the assessment activity on SPICA lead by ESA started in November 2007.

Both ESA and the SAFARI consortium have been making extensive studies during the assessment

phase, which is from Nov. 2007 to August 2009, and have been making good progress.


Date: 5/6/2023　


Calendar YearFiscal Year in J apanese Era

Fiscal Year FY2013 FY2014

20182007 2008 2009 2010 2011 2012 2013H26

2016 2017H29

2014 2015

FY2017FY2015 FY2016

J AXA

FY2007 FY2008 FY2009 FY2010 FY2011 FY2012

ESA

H27 H28H19 H20 H21 H22 H23 H24 H252019 2020 2021 2022

H34FY2018 FY2019 FY2020 FY2021 FY2022

H30 H31 H32 H33

ConceptStudiesPhase

ConceptDevelopment

Phase

ProjectFomulation

Phase

PreliminaryDesignPhase

FinalDesignPhase

Production and TestingPhase

LaunchOperations

Phase

InitialMission

OperationsPhase

NominalMission

OperationsPhase

AssessmentPhase (0/A)

DefinitionPhase (A/B1) Implementation Phase (B2,C/D)

Pre-Project Approval Review2008.5.12

Project Approval Review

Phase Ａ Phase B Phase C Phase D Phase EApproved

Launch

MDR2008.3.12

SRR SDR PDR CDR System AcceptanceReview LRR

CVSelection

CVDown Selection

CVFinal Selection

PDR CDR System AcceptanceReview

STA Delivery

Phase F

ExtendedMission

OperationsPhase

Fig.2.1 SPICA Schedule


Date: 5/6/2023　


3.　Summary　of　SPICA　Science　Objectives　

3.1　Overview

With SPICA, we aim to unveil one of the most fundamental questions: “the birth and evolution of

fundamental constituents of our Universe, such as our Solar system, our Galaxy, and the galaxies and

their large-scale structures.” In order to clearly define the SPICA mission requirements, we herewith

categorize the scientific objectives into three major subjects:

+ Resolution of Birth and Evolution of Galaxies

+ Revealing the Transmigration of Dust in the Universe

+ Thorough Understanding of Planetary System Formation

3.2　Resolution　of　Birth　and　Evolution　of　Galaxies

In this scientific subject, we define scientific targets as follows:

1. We will search for redshifted ionization lines (z>7) from low-metal objects (less than 10 -4) with

mid-IR spectroscopy, by which we intend to prove the existence of population III objects. We will

also investigate the formation of population III objects at z>3 through emission lines from hydrogen

molecules - important cooling lines of primeval molecular clouds using far-infrared spectrograph.

2. We will resolve the cosmic far-infrared background light into individual far-infrared objects with

spatial resolution more than 3 times higher than that of AKARI. We will then evaluate far-infrared

background fluctuations after removal of the individual objects, and reveal its origin through detailed

analysis such as multi-wavelength correlation.

3. We will reveal interstellar environment and dust emission characteristics of high-redshift

galaxies out to z~3 through PAH emission as well as atomic and molecular emission lines with

broad-band mid- and far-infrared moderate resolution spectroscopy. These observations allow us to

reveal the physical and chemical conditions of dusty galaxies in the early universe (up to 9 Gyr ago)

with precise correction for dust attenuation.

4. We will make infrared imaging and spectroscopic observations of TBD number of the forming

super-massive black holes (SMBHs),that cannot be observed easily in other methods due to the

obscuration of dust, from the present to the early universe. Supplementing these results with the

results of observations for the galaxy formation history, we will understand the role of SMBHs in the

galaxy evolution.


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5. In the early universe, where the star forming activities was at a peak, we will undertake wide-

area imaging survey, and observe the galaxy clusters and the large scale structures at infrared

wavelength, to which the redshifted emitting energy shifts. The large survey area (corresponding to

~300 Mpc) allow us to trace the large scale structures, and we will reveal the star formation history

in the early universe (up to 9 Gyr ago) as well as the mass assembly history and its environmental

effect on the galaxy evolution.

3.3　Revealing　the　Transmigration　of　Dust　in　the　Universe　


1. We will make some observations of several (>~5) dust-forming supernovae in nearby (<25Mpc)

galaxies within 1-2 years after the explosion. Changes in mid-infrared spectra of the supernova

during the processes, in which the dust is newly condensed in the SN ejecta gas and then it is cooled

down to the temperature of circumstellar pre-existing dust (~ a few hundred K), are examined to

specify its composition, size distribution and total mass.

2. We will make spatially well-resolved observations of faint dust shells around ~30 low- to

intermediate-mass evolved stars (e.g., AGB stars, planetary nebulae, novae etc.) in the Milky Way

and in the Magellanic clouds to investigate their mass-loss histories and the dust-formation

processes. Mid- to far-infrared spectra of spatially-resolved dust shell are used for constraining the

composition and the size distribution of dust condensed in the mass-loss gas.

3. We will make mid- to far-infrared spectroscopic observations of cold dense molecular clouds

with embedded young stellar objects in the Milky Way to detect the infrared bands of iron sulphide

grains and to demonstrate the link between the Glass with Embedded Metals and Sulfides (GEMS)

in Interplanetary Dust Particles (IDPs) and the interstellar grains. Then the grain growth scenario in

cold dense molecular clouds is explored.

4. We will make observations of about 30 SNRs so far detected in the infrared as well as those

detected in Objective 3. with imaging spectroscopy in the mid-to far-infrared to investigate the

composition/amount of formed dust, shock effects, and effects on the ISM (In total about 400 to 500

hours).

5. By mid- to far-infrared imaging spectroscopy (600 hrs in total), we will spectrally decompose

and spatially resolve emission from the ISM in 50 nearby galaxies of our AKARI sample, to track

galactic-scale material circulation from sources to sinks of the ISM in galaxies. The results also

complement the targets 1.~4.


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3.4　Thorough　Understanding　of　Planetary　System　Formation　


1. With the planet/star contrast ratio of 10-6 or better, we will directly detect gas exoplanets, and

perform spectroscopic observations of them to clarify the composition of the atmosphere.

Comparison with the results on our Solar System planets enables us to reveal the diversity of the

planetary systems.

2. With sensitive infrared spectroscopic observations, we will measure the gas in proto-planetary

disks, especially molecular hydrogen, and resolve the relation of gas mass with the age of primary

stars.

3. We will elucidate the geometric, physical and chemical structure of proto-planetary disks by

measuring the motion of gas with high-dispersion infrared spectroscopy.

4. With spatial resolution more than 3 times higher and sensitivity more than 10 times higher than

AKARI, we will detect a number of disks, which are comparable in amount of dust to our solar

system to understand the relationship with planetary systems observed in the other methods.

5. We will apply high-contrast IR coronagraphy to protoplanetary disks and debris disks, observe

their structures, and understand their relationship for disk evolution. Through infrared spectroscopic

observations with spatial resolution more than 3 times higher than AKARI, we will reveal

distribution and physical state of solid materials, particularly ice, in proto-planetary disks and dust

disks in the main-sequence stars.

6. With sensitivity more than 10 times higher than AKARI, we will make an unprecedented survey

of albedo, size, thermal inertia, and surface composition for primitive objects in the solar system.


Date: 5/6/2023　


4.　Mission　Requirements　

To achieve the scientific objectives described in chapter 3, diameter of telescope primary mirror

should be 3m class (current baseline: 3.5m) and its temperature should be lower than 6K. The core

wavelengths SPICA should cover are 5-210m. The wavefront error of the STA should be

minimized to achieve the diffraction limit at 5m. Sun – Earth L2 point is the optimum environment

to obtain excellent sky visibility and to cool the telescope. Onboard data generation rate is as high as

a few Mbps, therefore high-speed downlink (~10Mbps) to the ground station is necessary. In the

following subsection, the required specifications for the astronomical instrumentation at the focal

plane of STA (Focal-Plane Instruments: FPIs) and the required attitude control performance are

described.

4.1　Focal　Plane　Instrument　Specifications

To achieve the scientific objectives described in chapter 3, the following specifications for theFIP

are required.

Mid-infrared　camera　and　spectrometer

Requirement for the Imaging Specifications:

Wavelength coverage : 5～40μm

Spatial resolving power : diffraction limited

Field-of-View : 4’x4’ minimum, 6’x6’ is favorable

Imaging bands : multi-bands,~5, continuous coverage

Narrow band imaging (~100) is also required

Detection limit (1 hour, 5) : 10Jy at least, 1Jy is favorable

Requirement for the Mid-infrared Coronagraph Specifications:


Planet detection capability of nearly two-four times of diffraction-limited spatial

resolution

Contrast : 10-6

Low-resolution spectroscopy ( ~100)

Requirement for the Spectroscopic Specifications:

Low-resolution : ~100



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Continuum Sensitivity : 50μJy at least, 10μJy is favorable

Moderate-resolution : ~1000


Line sensitivity: 10-19W/m2 at least


High-resolution : ~30,000 (velocity resolution of 10km/s)

Wavelength coverage <20μm

Far-infrared　camera　and　spectrometer

Requirement for the Imaging Specifications:

Wavelength coverage : 35～210μm baseline

Spatial resolution : diffraction limited

Field-of-View : 2’x2’ minimum @40-70 m deep imaging

Imaging bands : multi-bands,~5, continuous coverage

Detection limit (1 hour, 5) : 100Jy @50m at least, 50Jy is favorable (source

confusion limit at wavelength >70m)

Requirement for the Spectroscopic Specifications:

Low-resolution : ~100

Wavelength coverage : 35～210μm at least

Continuum Sensitivity : 1mJy @50m at least, 0.2mJy is favorable

Moderate-resolution : ~2000 at 100μm

Wavelength coverage : 40～200μm (extension up to ~400m is favorable)

Line sensitivity: 10-18W/m2 at least


Coronagraph　Instruments

Wavelength coverage : 3.5～27μm (shorter wavelength is optional)

Coronagraphic method : Binary-shaped pupil mask (baseline)

Contrast : 10-6

Inner working angle (IWA) : ~ 3.3 Dbinary-shaped pupil mask）

Outer working angle (OWA) : ≤16 D


Date: 5/6/2023　


Detector : Si:As 1K×1K (InSb detector is optional)

Field of view : 1.0×1.0 [arcmin2]

Spectral resolution : R=20~2004.　2　Pointing　Control　RequirementThis section describes the pointing control requirements in each of the mission operation modes

below. The requirements are applicable to all the FPIs unless otherwise specified, and the number is

specified by the most stringent requirements.

(1)　Pointing　Mode

In this mode, one of the FPI performs its observation. The requirements are listed below.

RequirementsObservation Mode

Other than Coronagraph Coronagraph

Pointing Control Accuracy 0.135 [arcsec](3σ) 0.03 [arcsec](3σ)

Pointing Stability 0.075 [arcsec](0-P)/200sec 0.03 [arcsec](0-P)/20min

Table 4.1 Pointing Control Requirements

Note: 0-P = 0 to peak (half the width of pointing fluctuation over a given period of time)

(2)　Step　Mode

For slight change of the pointing direction, spacecraft attitude is to be shifted as specified in the

table below. The requirement is not applicable to either BLISS or coronagraph observation.


SAFARI Others

Step Angle 108 [arcsec] 0.075 to 22.8 [arcsec]

Step Angle Accuracy +/- 12 [arcsec] +/- 0.075 [arcsec]

Step Direction every direction every direction

Settling Time within 100 seconds within 100 seconds

Table 4.2 Step Mode Requirements

(3)　Non-sidereal　Tracking　Mode

In order to track a non-stationary object, spacecraft attitude is to be moved as specified in the table

below. In the mode, pointing control accuracy is supposed to meet the requirements described in (1).

Requirements Description


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Tracking Speed Range up to 10 arc-seconds /minute

Tracking Duration up to 1200 seconds

Table 4.3 Non-sidereal Tracking Requirements

(4)　Slow-scan　Mode

In this mode, a stationary target star is measured by continuously shifting the spacecraft attitude

with constant speed and direction as specified in the table below. The requirement is not applicable

to either BLISS or coronagraph observation.


SAFARI Others

Slow-scan Speed Range 10 to 72 [arcsec/s] 0.054 to 2.28 [arcsec/s]

Slow-scan Speed Accuracy less than 1% less than 10%

Slow-scan Duration up to 600 [s] 5 to 50 [s]

Table4.4 Slow-scan Requirements

(5)　Pointing　Reconstruction　Mode

The requirement is applicable only to coronagraph observation prior to performing the pointing

mode. In this mode, spacecraft attitude is controlled so that the Coronagraph Focal Plane Camera (C-

FPC) can acquire a target star within its field of view.

(6)　Attitude　Maneuver　Mode

In order to change the observing direction, attitude maneuver is to be performed so that the

maximum maneuver angle of 180 degrees can be realized within 30 minutes. In this mode, mission

observation cannot be performed. In addition to the initial and final pointing directions, mid-course

direction is supposed to also satisfy the sun avoidance angle.


Date: 5/6/2023　



Date: 5/6/2023　


5.　Mission　Profile　

SPICA is the JAXA’s future science program and is planned to be launched in 2018. The mission

follows on the highly successful AKARI mission both scientifically and technically. High

photometric sensitivity in observations in mid- and far-infrared are realized by the 3m class

telescope (current baseline: 3.5m), which is actively cooled to below 6 K (target:<6K) to eliminate

the non-astronomical photon noise effectively. High spatial resolution is achieved thanks to the large

aperture, monolithic primary mirror and the appropriate tolerances on the telescope and mirror

surfaces, which are designed to have diffraction-limited performance at 5μm.

The envisaged launcher for SPICA is the Japanese H-IIB rocket. The rocket, which is able to

deliver more than 4,000 kg load to L2 transfer orbit, launches SPICA from the JAXA Tanegashima

Space Center. H-IIB is an upgraded version of the H-IIA with the aim of offering a new possibility

for future missions, including cargo transport to the International Space Station (ISS) and to the

Moon. H-IIB launch vehicle is a two-stage rocket using liquid oxygen and liquid hydrogen as

propellant and has four strap-on solid rocket boosters (SRB-A) powered by polybutadiene. The H-

IIB has two liquid rocket engines (LE-7A) in the first-stage, whereas the H-IIA has one. It has four

SRB-As attached to the body, while the standard version of H-IIA has two. In addition, diameter and

total length of the H-IIB's first-stage body expanded. H-IIB’s diameter (5.2 m) is 1.2 m longer and

its total length is 1m longer than H-IIA’s counterparts, respectively. The H-II Transfer Vehicle to be

launched with H-IIB in 2009. The payload fairing of the launcher has a diameter of 4600 mm,

(HTV) is planned which accommodates a 3m-class telescope and associated thermal baffling. The

satellite will be launched at ambient temperature, and cooled down below 6K (target: <5K), which is

the operating temperature during the Checkout Phases. This approach brings several benefits, the

greatest of which is realization of the lighter and simpler spacecraft.


Date: 5/6/2023　


Fig.5.1

The envisaged orbit for SPICA is a Halo orbit around the Sun – Earth Lagrange Point 2 (S-E L2).

The S-E L2 provides SPICA with a benign and stable thermal environment, which is required to cool

the payload lower than 6 K, as well as a good instantaneous sky visibility. The S-E L2 Halo orbit is

an elliptical orbit with a period of about 120 days.


Date: 5/6/2023　


100

50

0

-50

-1000 50 100 150 200

[×10000km]

[×10000km]z

x

100

50

0

-50

-10050 100 150 200

[×10000km]

[×10000km]

Geocentric Rotational Coordinate(xy: Ecliptic Plane)

y

x

Geocentric Rotational Coordinate

(xy: Ecliptic Plane)

L2

L2

100

50

0

-50

-100-100 -50 0 50 100

[×10000km]

[×10000km]z

y



L2

20 day

40

80

Earth

100

60

0

Halo Orbit

HOI

Earth

20 day40

80

100

60

HOI

Halo Orbit Halo Orbit20 day

40

80

100

60

HOI

Earth

The guaranteed lifetime of SPICA is 3 years, with the goal of a 5 year-extended operation. The

absence of cryogens onboard allows to extend the nominal lifetime beyond the nominal duration

(finally limited by the AOCS propellant and onboard failures). 4K Mechanical Cooler remains fully

operative until the cooling power degrades below 40mW (Nominal cooling power is 50mW).

There are 4 operation phases in the period between the satellite installation into the launch facility

and the end of the mission.

射場整備

(Launch Operations Phase)初期運用フェーズ

(Initial Mission Operations Phase)定常運用フェーズ

(Nominal Mission Operations Phase)後期運用フェーズ

(Extended Mission Operations Phase)

追跡管制隊発足

打上げTBD時間前(L-TBD Hours)

打上げ(Launch)

クリティカル運用終了(End of Launch Operation )

観測軌道投入完了(L2 Halo insertion complete)

打上げ3年後(L+3Years)

停波(Termination)

Fig.5.3

Fig.5.2


Date: 5/6/2023　


Launch operations Phase is a period from the satellite arrival at the Tanegashima Space station to

the end of the launch operation. Initial Mission Operations Phase starts at the end of the launch

operations phase. In this phase, the satellite is injected into the observation orbit, the S-E L2 Halo.

During the injection, function checkout and test observations are performed. 18 hours after the

launch, first orbit maneuver is operated. After that, a few small orbit maneuvers are conducted and

20 to 40 days after the launch, the satellite reaches near the S-E L2 point. In parallel, the telescope

starts to be cooled. 120 days after the launch, the satellite is injected into the observation orbit. At

this moment, we assume that it takes 168 days to cool the telescope with using 2 redundant 4K

Mechanical Cooler at the same time.

100

50

0

-50

-10050 100 150 200 250

[×10000km]

[×10000km]

Az = 30万km

z

x

100

50

0

-50

-10050 100 150 200 250

[×10000km]

[×10000km]



Az = 30万km

y

x



L2

L2

100

50

0

-50

-100-100 -50 0 50 100

[×10000km]

[×10000km]z

y



L2

0 day

120

90

30 60

150

0 day

120

90

30

60

150

Az = 30万km

Nominal Mission Operations Phase is the planned operation period and supposed to be 3 years after

Fig.5.4


Date: 5/6/2023　


the launch. Observation is performed with only one focal plane instrument working at each operation

scenario in order to keep the telescope cooler than 6K (target: <5K). The high gain antenna with 2-

axis gimbal is able to send mission data while telescope is observing the space. Extended Mission

Operations Phase is a period from the end of Nominal Mission Operations Phase to the termination

of the satellite. Operation time is expected to be more than 2 years.

Since SPICA is the unprecedentedly large satellite mission, its operations such as observation and

data distribution cannot be run only by the volunteer scientists. Thus, JAXA makes an operation

framework supported by dedicated staff members, while taking advantage of scientists’ active

involvement in the project operations.

For the purpose, we shall set up the following two sections in the Tracking System.

1. Mission Operation Center (MOC)

This section, which is set up in JAXA, is responsible for operations such as confirming the project

soundness, and establishing the contingency plan.

The basic functions are:

- generating and transmitting commands

- receiving telemetry data

2.Science Operation Center (SOC)

This section, which is set up in JAXA, and in the countries concerned if needed, is responsible for

the following two functions:

a. observation operations.

-supporting users by offering information and tools to make observation proposals

-compiling observation proposals by gathering them through a public subscription

-supporting TAC (Telescope Allocation Committee), which is responsible for judging scientific

values of the proposals, and allocating the observation time for the proposals

-establishing the observation plan on the basis of the accepted proposal and offering the plan to

SOC

b. data analysis and archiving

-processing scientific data with support from the observation instruments development team

-archiving the data with support from other data-related sections in JAXA

-distributing the data to the observers with critical information such as telescope direction, and

analysis tools

-publishing the data with support from other data-related sections in JAXA.


Date: 5/6/2023　


The mission requires eight hour ground contact per day at minimum. The spacecraft and

instruments are supposed to have sufficient autonomic functions to be tolerant to the long-time

disconnection from ground contact. The mission will be greatly enhanced by additional ground

contact time, which leads to increase in the telemetry budget.

SPICA Tracking System

Mission Operation Center（MOC)

J AXA Ground Network(Usuda, Uchinoura)

Science Operation Center（SOC)

Science Data Center（SDC)

ESA Ground Network(Cebreros)

Command / Telemetry

Operation Plan

J AXA

ESA KOREA USA

Mission Data

SPICA Scientist

Mission Data/Analysis tool

Fig. 5.4Fig.5.5


Date: 5/6/2023　


6.　Spacecraft　Design　

6.1　Overview

6.1.1　Layout

The layout of the SPICA spacecraft is schematically illustrated in Figure 6.1. The SPICA consists

of the Payload Module (PLM) and the Bus Module (BM). The Cryogenic Assembly of PLM cools

the Science Instrument Assembly (SIA), which includes the SPICA Telescope Assembly (STA) and

the Focal-Plane-Instrument Assembly (FPIA), with a mechanical cooler system and a passive

radiative cooling system. The PLM is connected to BM through a truss structure. The Bus Module

prepares the capabilities of electric power supply, communications, attitude control, and so on. The

subsystems of SPICA are summarized in Figure 6.2. Major specifications of the BM subsystems are

described in following sections in this chapter. The description of the PLM is given in chapter 7.

Fig. 6.1 Layout of the SPICA spacecraft; (a) stowed in the rocket fairing and (b) 3D cut-away view

illustrating the thermal stages.


Date: 5/6/2023　


Fig. 6.2 SPICA subsystems

Payload Module (PLM)

Cryogenic Assembly (CRYO)

Scientific Instrument Assembly (SIA)

SPICA Telescope Assembly (STA)

Mechanical Cooling System (MCS)

Focal-Plane-Instrument Assembly (FPIA)

CRYO Driving & Monitoring System (CRYO-DMS)

CRYO Driving & Monitoring System (CRYO-DMS)

Thermal Insulation & Radiative Cooling System (TIRCS)

FPI Harness

SIA Warm Electronics

Bus Module(BM)

Bus Structure Subsystem (STR)

Bus Thermal Control Subsystem (TCS)

Electric Power Subsystem (EPS)

Communications Subsystem (COM)

Data Handling Subsystem (DHS)

Attitude & Orbit Control Subsystem

(AOCS)

Reaction Control Subsystem (RCS)

Wire Harness Subsystem (WHS)


Date: 5/6/2023　


6.1.2　Satellite　Mass　and　Electric　Power

The mass estimations of SPICA subsystems are summarized in Table 6.1. The total wet mass is

approximately 4,000kg.


Date: 5/6/2023　


Payload Module Mass (including maturity margins)

Cryogenic Assembly 1200 Scientific Instrument Assembly Telescope Assembly 700 Focal-Plane-Instrument Assembly * 200 SIA Warm Electronics 100 Cryogenic Electronics 90 Sub Total 2290 Payload Module Total (including maturity margins) 2290 Bus Module Bus Structure 330.0 Bus Thermal Control 66.0 Electric Power Power Control Unit / Shunt / Battery 31.9 Solar Array Paddle 93.5 Communications 115.5 Data Handling 22.0 Attitude & Orbit Control 105.6 Reaction Control 110.0 Wire Harness 38.5 Sub Total 913 Bus Module Total (including maturity margins) 913 System Mass (kg) Payload Module Total (with margin) 2290 Bus Module Total (with margin) 913 Subtotal (with margin) 3203 System level margin (Subtotal x 20% ) 641 Propellant 220 Total (with maturity margin and system margin) 4064

Table 6.1 Mass Estimation (Aug. 28, 2009)

*excluding BLISS

Electric power required by each subsystem is given in Table 6.2. This table shows the maximum

electric power in the observation phase.


Date: 5/6/2023　


Subsystem

Elec. Power

(W, Max)

Payload Module

Cryogenic Assembly 900

Scientific Instrument Assembly* 235

Bus Module

Bus Structure 0

Bus Thermal Control 160

Electric Power 35

Communications 183

Data Handling 100

Attitude & Orbit Control 130

Reaction Control 185

Total 1,928

Margin (20%) 386

TOTAL (with margin) 2314

Table 6.2 Electric Power Requirements

*excluding BLISS


Date: 5/6/2023　


6.2　Mechanical　Structure

Figure 6.3 shows SPICA on-orbit configuration, which consists of a Bus Module, a Payload

Module, two Solar Array Paddles (SAP) and a High Gain Antenna (HGA). The two-axis gimbal

maximum dimension of SPICA is about 4600mm (width) X 7500mm (height).

Fig. 6.3 SPICA on-orbit configuration

To accommodate the size of a Japanese H-IIB launch vehicle, a 5S-H fairing and a 2360SA PAF are

the current candidates.

H-IIB is planned to be launched in 2009 for the first flight. Mechanical environment of H-IIB is

estimated to be the same as that of H-IIA204 launch vehicle, which was previously the candidate

launcher.

5S-H faring is the largest faring whose envelope for a satellite is about Ф4600mm X 8000 (height

Payload Module

Bus Module

SAP


Date: 5/6/2023　


of the cylindrical part).

2360SA PAF is made of aluminum alloy and attached to a satellite at 8 interface points, which are

separated by separation nuts. The advantage of the PAF is that it is lower at height than others.

Figure6.4 shows a launch configuration, in which SAPs and HGA are stored in the Bus Module.

　　　　　　　　　Fig. 6.4 SPICA launch configuration

H-IIB interface requires that stiffness of SPICA satellite shall satisfy the following condition:

longitudinal direction >= 30Hz (6DOF at interface points are fixed)

lateral direction >= 10Hz (6DOF at interface points are fixed)

Stiffness requirement for STA is assigned to be tentatively

longitudinal direction >= 60Hz (6DOF at interface points are fixed)

lateral direction >= 30Hz (6DOF at interface points are fixed)

torsion around longitudinal axis >= 30Hz (6DOF at interface points are fixed)

Assuming the above stiffness and mass properties of Payload Module, environmental conditions for

STA is defined in a document JAXA_SPICA_IF0002 “SPICA ENVIRONMENTAL CONDITIONS


Date: 5/6/2023　


FOR SPICATELESCOPE ASSEMBLY (STA)”.

A set of truss and panel structure has been chosen as the Bus Module main structure because of the

simple load path as well as the flexibility of side panel design and integration. As the number of PAF

interface points is eight, there are also eight mechanical interface points at Ф3500mm circle on the

upper panel of the Bus Module. The load from the heavy Payload Module is transferred to Bus

Module truss and to the launcher efficiently through these interface points. On the lower panel, RCS

subsystem is installed and side panels are capable of installing all the other electrical components

including mission electronics and drivers of cryogenic coolers.

Total mass of structure subsystem is calculated to be about 300kg.

Although TRL of these structures is high enough, following subjects are to be addressed in the

process of structural design.

- Vibrational and Quai-static load estimation by coupling mathematical model of Payload Module

and Bus Module

- Thermo-elastic analysis to estimate the effect of thermal distortion of Bus Module structure to

WFE of the primary mirror

- Effect of fill factor to acoustic environment

- Possible method for reducing micro vibration from cryogenic compressor, reaction wheel and so on

6.3Thermal　Control

In this section, the thermal control of the Bus Module is described. The thermal control of the PLM

is described in section 7.1

6.3.1　Thermal　Environment

During the early stage after the launch, a spacecraft is exposed to infrared and albedo from the

Earth as well as direct sunlight. When SPICA is on a transfer orbit to L2 and L2 orbit, direct

sunlight is the only environmental heating. The variation of the thermal environment is relatively

moderate.

6.3.2　Thermal　Design

The functional requirement is the thermal control of the Bus Module.


Date: 5/6/2023　


The performance requirement is to maintain all the items and the interfaces between the Payload

Module and the Bus Module within their allowed temperature limits during all mission phases.

Fig.6.5 shows the schematic of the structural panels. The following is the thermal condition of these

panels in a typical observation mode. The sunlight always illuminates the side panel (2), (3) and (4).

Since the side panels (2) and (4) thermally are coupled with the solar arrays, they are inappropriate

as radiators. The other panels are appropriate to reject heat since there is little heat backloads on

them from environment and the solar panels.

6.3.3　Thermal　Conditions　at　the　Interface　with　Payload　Module

Fig.6.6 shows the thermal conditions at the interface with the Payload module. The upper limit

temperature and the heat flow limits of interfaces are tentatively given. The conductive heat

exchange is assumed to be 0W in the Bus Module thermal design.

The side panels are thermally isolated from truss structures. They load the equipments of the Bus

Module and reject their waste heats. The lower panel is a dedicated radiator for the upper panel and

the truss structures.

The very preliminary thermal analysis shows that there is a solution for a typical observation case.

Sunlight

Upper Panel

Lower Panel

Side Panel (5)

Side Panel (1)

Side Panel (2)

Side Panel (3)Side Panel (4)

Side Panel (6)

Side Panel (7)

Side Panel (8)

Fig.6.5 Schematic of the structural panels


Date: 5/6/2023　


6.4　Electrical　Power　Subsystem

6.4.1　Overview

The electrical power subsystem (EPS) provides required electrical power to the bus and the mission

instruments through the mission life. The major components of the EPS are shown in Table 6.3. It is

equipped with the solar array paddle (SAP), the power control unit (PCU), the shunt dissipater

(SHNT), and the battery (BAT). They are free from single point failures in principle.


Date: 5/6/2023　


Equipments　on　EPS Quantity Weight

Solar Array Paddle (SAP) 2 84.9 kg

Power Control Unit (PCU) 1 9.4 kg

Shunt Dissipater (SHNT) 1 8.0 kg

Battery (BAT) 1 11.7 kg

Total 114.0 kg

Table 6.3 EPS Configuration

6.4.2　Solar　Array　Paddle　(SAP)

The SAP generates the electrical power for the satellite from sunlight. The solar array power

generation is more than 2.4 kW (EOL) in the S-E L2 Halo orbit.

The SAP configuration is listed in Table 6.4. It comprises two wings, each consisting of three solar

panels, a deployment hinge, a holding mechanism, and a yoke. The solar panels are equipped with

triple junction solar cells.

The wings are stowed with the panels wrapped around the satellite body on launch, and deployed

over the sun shield after the satellite is separated from the rocket.

Equipments on SAP Quantity Weight (2 Wing Total)

Solar Panels 3 × 2 wings 63.1 kg

Holding Mechanism 2 8.4 kg

Yoke 2 4.8 kg

Hinge 2 8.6 kg

total 84.9 kg

Table 6.4 SAP Configuration

6.4.3　Shunt　Dissipater　(SHNT)

A 50-V unregulated bus shall be adopted. The SHNT diverts excess power to regulate the bus

voltage when the power generated by the solar array exceeds the load power.


Date: 5/6/2023　


A digital shunt is employed for lower weight and heat generation compared to analog shunts.

6.4.4　Power　Control　Unit　(PCU)

The PCU provides interfaces to the telemetry, controls the charge and discharge of the battery, and

distributes the bus voltage to the load. It protects the battery by stopping charge in case of over

voltage (4.2 V/cell) and over temperature (15ºC or 25ºC), and turns off low-priority equipments in

case of under voltage (3.0 V/cell or 2.5 V/cell).

6.4.5　Battery　(BAT)

The BAT provides electricity during the launch, orbit controls, and attitude deviations; each

operation requires 400 Wh or more. It comprises eleven 23.5-Ah class Li-ion battery cells in series,

protection circuits, and voltage balancing circuit. One cell of short-circuit or open-circuit failure is

admissible.

The number of charging and discharging cycles is 50 or less on the ground and 100 or less in space.

The aged deterioration rate is 3 % or less on the ground and 2 % or less in space.


Date: 5/6/2023　


6.4.6　Performance

The specification performance of EPS is listed in Table 6.5.

Item Performance Remarks

Bus Voltage 32.5 V – 52.0 V Ripple Noise ≤ 500 mVp-p

Power Delivery Capability 2.4kW

SAP Generated Power 2.4 kW 2.4 kW(EOL), 2.7 kW(BOL)

Incident deg. -30 to +30 deg.

Temperature 100 ºC

Output Voltage 55 V

Natural Vibration Freq. 1 Hz or more during deployment

SHNT Shunted Power 3.2 kW, 20 levels 160 W/level, 20 levels in total

BAT Capacity 23.5 Ah

Configuration 11 S, 1 String

Voltage Range of a Cell 3.0 V – 4.1 V

Charge Bypass Voltage 4.1 V/cell Linear Shunt

Over Voltage Protection 4.2 V/cell

Under Voltage Control 3.3 V or 2.75 V 1 Cell

Max. DOD 45.0 % BOL

PCU Battery Charging Control 0.5 A/CC

34 V – 47.5V/CV

CV in 256steps

Battery Voltage Monitoring 0 – 5 V/cell 10 bit

Over Voltage Protection for

Battery

4.2 V / Cell

Over Temperature Protection

for Battery

15 deg. or 25 deg.

Under Voltage Control for

the Battery

3.3 V or 2.75 V/cell


Date: 5/6/2023　


Table 6.5 Performance of EPS

6.5　Communications　Subsystem

The Communications Subsystem (COM) onboard SPICA provide uplink command, turnaround

ranging, and downlink telemetry capabilities. Block diagram of dual S-band up /S- and X-band down

telecommunication systems is shown in Fig. 6.7. This design enables us to provide high throughput

data transmission without any interruption caused by the turnaround ranging and low-rate HK

telemetry transmission from the S-E L2 Halo orbit.

The component list of the onboard system is shown in Table 6.6. The total weight and power

consumption are 104.96 kg and 182.5W, respectively.

S-band and X-band communication link availability is as shown in Table 6.7 and 6.8. The SPICA

mission requires high-speed telemetry downlink, 11Mb/s using HGA and 1Mb/s using MGA.

Combination of the UDSC 64m and USC 34m receiving antennae met the SPICA TT&C

requirements.

Fig. 6.7 Block diagram of dual S-band up /S- and X-band down onboard telecommunication

system.

X-band high speed

Telemetry downlink

S-band low speed

TT&C link


Date: 5/6/2023　



Date: 5/6/2023　


units name 　components namemass　

（kg）

power dissipation (W)

XHGA X-Band High-Gain Antenna 61.435

XHGA Control Unit 　 11.3

XMGA X-Band Medium-Gain Antenna2.61

XMGA Control Unit 　 6

XSW-1X-Band Switch (DPDT, High Power)

0.2

XTWT-1 X-Band Travelling Wave Tube 2.34 56.2

XEPC-1X-Band Electronic Power Conditioner

XTWT-2 X-Band Travelling Wave Tube 2.34 0

XEPC-2X-Band Electronic Power Conditioner

XHYB-1 X-Band Hybrid Divider 0.07

XTX-1 X-Band QPSK Transmitter 2.75 12.5

XTX-2 X-Band QPSK Transmitter 2.75 12.5

SLGA-1 S-Band Low-Gain Antenna 0.7

SLGA-2 S-Band Low-Gain Antenna 0.7

SDIP-1 S-Band Diplexer 1.7

SDIP-2 S-Band Diplexer 1.7

SSW-1S-Band Switch (SPDT, High Power)

0.2

SSW-2S-Band Switch (SPDT, High Power)

0.2

STRP-1(with SSSPA-1) S-Band Transponder 7 50

STRP-2(with SSSPA-2) S-Band Transponder 7 10.3

　total 104.96 182.5

Table 6.6 Components list of the SPICA onboard TT&C systems


Date: 5/6/2023　


S-band Uplink USC 34m (Command 1Kb/s) S-band Down-link USC34m (Telemetry 2kb/s)

Frequency MHz 2100 Frequency MHz 2250

Transmitted Power dBm 70 10kW Transmitted Power dBm 40 10W

Feeder Loss dB -0.5 Feeder Loss dB -

Antenna Gain dBi 55.6 KSC 34m Antenna Gain dBi 0

Pointing Loss dB - Pointing Loss dB -

EIRP dBm 125.1 EIRP dBm 40

Path Loss dB 224.8 2000000km Path Loss dB 225.4

Absorption Loss dB -0.3 Absorption Loss dB -0.3

Rain Attenuation dB 0.0 Rain Attenuation dB 0.0

Polarization Mismatch dB - Polarization Mismatch dB -


Antenna Gain dBi 0 LGA Antenna Gain dBi 56.9 UDSC34m

Feeder Loss dB - Feeder Loss dB -

Received Power dBm -100 Received Power dBm -128.8

System Noise Temp. dBK 28 627.1K USB-TRP System Noise Temp. dBK 19.8 95.9K

Boltzmann Const. dBm/Hz/K -198.6 Boltzmann Const. dBm/Hz/K -198.6

Noise Power Density dB/Hz -170.6 Noise Power Density dB/Hz -178.8

C/No dBHz 70.6 C/No dBHz 50

Carrier Command 1kb/s Carrier Telemetry 2kb/s

Required Bandwidth dBHz 30 (1kHz) 30 Required Bandwidth dBHz 30 33.1

Filter Bandwidth dBHz - - Filter Bandwidth dBHz - -

Modulation Index Rad - 0.4 Modulation Index Rad - 1.1

Modulation Loss dB 1.2 11.9 Modulation Loss dB 3.7 4.1

Hardware Loss dB - - Hardware Loss dB - 1.5

Coding Gain dB - - Coding Gain dB - 6.5

Required C/No dB - - Required C/No dB - -

Required S/No dBHz 13.5 15 Required S/No dBHz 13.5 9.6

Required C/No dBHz 44.7 56.9 Required C/No dBHz 47.2 41.8

Margin dB 25.9 13.7 Margin dB 2.8 8.2

Table 6.7 S-band communication link availability


Date: 5/6/2023　


X-band Down-link USC 34m (11Mb/s) X-band Down-link UDSC 64m(1Mb/s)

Frequency MHz 8500 Frequency MHz 8500

Transmitted Power dBm 43 Transmitted Power dBm 43

Feeder Loss dB 2.0 Feeder Loss dB 2.0

Antenna Gain dBi 40.8 D=1.6m, μ=0.6 Antenna Gain dBi 17.6


EIRP dBm 81.8 EIRP dBm 58.6

Path Loss dB 237.0 2000000km Path Loss dB 237.0

Absorption Loss dB 0.4 Absorption Loss dB 0.4

Rain Attenuation dB 2.5 5mm/h Rain Attenuation dB 2.5

Polarization Mismatch dB - Polarization Mismatch dB -


Antenna Gain dBi 67 KSC 34m Antenna Gain dBi 72.5 UDSC64m

Feeder Loss dB - Feeder Loss dB -

Received Power dBm -91.1 Received Power dBm -108.8

System Noise Temp. dBK 23.1 205.6K System Noise Temp. dBK 18.7 74K

Boltzmann Const. dBm/Hz/K 198.6 Boltzmann Const. dBm/Hz/K 198.6

Noise Power Density dB/Hz -175.5 Noise Power Density dB/Hz -175.5

C/No dBHz 84.4 C/No dBHz 71.1

Telemetry 11Mb/s Telemetry 1Mb/s

Required Bandwidth dBHz 70.4 Required Bandwidth dBHz 60

Filter Bandwidth dBHz - Filter Bandwidth dBHz -

Modulation Index Rad - Modulation Index Rad -

Modulation Loss dB - Modulation Loss dB -

Hardware Loss dB -2.0 Hardware Loss dB -2.0

Coding Gain dB 4.5 Coding Gain dB 4.5

Required C/No dB - Required C/No dB -

Required S/No dBHz 11.3 Required S/No dBHz 11.3

Required C/No dBHz 79.2 Required C/No dBHz 68.8

Margin dB 5.2 Margin dB 2.3

Table 6.8 X-band communication link availability


Date: 5/6/2023　


6.6　Data　Handling

This section presents the overview of the data handling subsystem of SPICA.

6.6.1　Architecture　

The data handling subsystem and other subsystems that process data will be developed according to

the standard communications and data handling architecture developed by JAXA/ISAS for its

science projects. The purpose of using this architecture is to promote sharing of onboard components

among different projects.

This architecture has three sub-architectures: physical architecture, functional architecture, and

protocol architecture.

The physical architecture defines physical elements and how to connect them physically. There are

two types of physical elements: intelligent nodes and non-intelligent nodes. Intelligent nodes are

physical elements having one or more processors, while non-intelligent nodes are those without a

processor. For intelligent nodes, onboard computers that conform to the SpaceCube architecture

should be used. The SpaceCube architecture specifies the requirements for the operating system,

interfaces and middleware to be used on intelligent nodes. For non-intelligent nodes, it is

recommended that standard interface cards be used as the interface to the network. Nodes are

connected with SpaceWire networks, the details of which are specified by the protocol architecture.

The functional architecture defines functional elements and how to connect them functionally. A

non-intelligent node should be monitored and controlled by an intelligent node directly connected to

the non-intelligent node. An intelligent node should be monitored and controlled by another

intelligent node that is directly connected to the intelligent node and closer to the communications

subsystem than the node. For monitoring and controlling the function of nodes, the Spacecraft

Monitor and Control Protocol should be used.

The protocol architecture specifies the communications protocols that should be used between

physical components and between functional components. Between physical components, the

SpaceWire protocol, the Remote Memory Access Protocol, and SpaceWire RT should be used.

Between functional components, the Spacecraft Monitor and Control Protocol and the Space Packet

Protocol should be used.


Date: 5/6/2023　


6.6.2　Functions

The functions of the data handling system are as follows.

(1) Command processing

Commands are received by the data handling subsystem from the ground through the

communications subsystem as a sequence of Space Packets, each contained in a Transfer Frame.

Received command Packets are distributed to the other components based on the ID contained in the

header of each Packet. The data handling subsystem also receives from the ground a timeline, which

is a sequence of time-tagged commands. The received timeline is stored in the data handling system

and individual commands contained in it are distributed to the destination components according to

the time tags.

(2) Telemetry processing

Telemetry is collected by the data handling subsystem from the other components as sequences of

Space Packets. Collected Packet are transmitted to the ground through the communications

subsystem in Transfer Frames and/or stored in the onboard data recorder. The recorder has multiple

partitions and each Packet is stored in the partition specified by a value in its header. The contents of

the data recorder are sent to the ground during a tracking pass and the partition with the highest

priority is read out first.

(3) Autonomous control

SPICA has an autonomy function of operation. The function (1) is to be used mainly for

contingency cases of satellite operation and (2) is not meant to be used for routine observations. The

SPICA's autonomy function is controlled by the main data processor CDHU (Central Data Handling

Unit) on the basis of the telemetry data. We are not planning to use special information not

downlinked to the ground for the autonomy function, and each data processor, except for CDHU and

AOCS, is not supposed to have autonomy functions.

The most representative case of the contingencies in the autonomy function is the UVC (Under

Voltage Control). Most of the contingency cases related with attitude control are to be handled by

AOCU.


Date: 5/6/2023　


The current baseline for routine observations is that each of them is not supposed to use the

autonomy functions. Every operation procedure is to be uploaded explicitly from the ground.

The data handling subsystem checks the values of the specified telemetry items periodically, and

when a value or a set of values violates the predefined rules, it issues a command specified by the

rule. It issues a command when an alert message is received from a component, too. It may also

generate a command when the expected result is not obtained as the result of execution of the

timeline mentioned in (1) above.

(4) Time management

The data handling subsystem manages the onboard master clock and distributes its value to the

other components through the onboard networks at a predefined interval. It also generates special

telemetry Packets in order for the ground system to correlate the onboard clock with the universal

time.

6.6.3　Performance

The following table shows the tentative performance of the data handling subsystem.

Processor Processor type HR 5000, 33MHz

Operating System T-kernel

Ground interface Command rate 1000sps max

Telemetry rate 13Msps (mission), 2ksps (HK)

Onboard interface Transmission rate TBD

Data recorder Capacity 48Gbytes min

Recording rate 4Mbps min

General Mass 20kg

Power 100W

Table 6.9 Performance of the data handling subsystem (tentative)

Note: sps=symbols per second

6.6.4.　Baseline　configuration


Date: 5/6/2023　


The following figure shows the baseline configuration of the data handling subsystem.

Fig.6.8 Baseline configuration of data handling subsystem

6.7　Attitude　and　Orbit　Control　System　(AOCS)

This section describes SPICA Attitude and Orbit Control System (AOCS) together with its related

systems that realize fine pointing performance.

(1)　AOCS　Block　Diagram

Fig.6.9 depicts AOCS Block diagram. AOCS consists of attitude sensors (Inertial Reference Unit

(IRU), Star Trackers (STT), Fine Sun Sensors (FSS), Magnetometer (MAG), Accelerometer (ACC)),

processors (AOCP), and actuators (Reaction Wheels (RW), Drive Unit (DRU)). For fine pointing,

the IRU and the STTs are applied to the fine attitude determination system (ADS), and for fine

attitude control the RWs are applied.

However, the primary AOCS is not good enough to meet the mission pointing requirement. Figure

6.9 also describes instruments mounted on the Instrument Optical Bench (IOB). For mission

observation modes other than Coronagraph observation, Focal Plane Camera (FPC-G) is utilized as

fine guidance sensor that can reduce alignment error and random noise than the STTs, thus improve

the ADS accuracy. For coronagraph observation, another Focal Plane Camera (C-FPC), which is

built in the coronagraph, and Tip-Tilt Mirror (C-TTM) are utilized to further improve the pointing

accuracy by controlling the TTM according to the C-FPC information. In this mode, AOCS need to


Date: 5/6/2023　


provide conditions required for C-FPC acquisition and C-TTM unloading.

Attitude andOrbit controlProcessor

Inertial Reference Unit

Magnetometer

Accelerometer

Focal Plane Camera

Coronagraph

Drive Unit RCSX-ANT

Star Tracker

Fine Sun Sensor

Star TrackerStar Tracker

Fine Sun Sensor

AccelerometerAccelerometer

Reaction WheelReaction WheelReaction WheelReaction Wheel

Attitude andOrbit controlProcessor

Focal Plane Camera

Tip-Tilt Mirror Controller

IOB Instruments

ActuatorsSensors Processor

Fig.6.9 AOCS Block Diagram

(2)　AOCS　Functions

The AOCS main functions are listed in Table 6.10. Mission observation is realized by the function

defined by “Attitude Control for Observation”.

Functions Descriptions

Attitude Acquisition After the separation from the launch vehicle or after the loss of attitude

by anomaly, rate damping and sun acquisition are performed.

Safe Attitude Control After attitude acquisition, sun pointing attitude is performed to acquire

solar power and communication link.

Attitude Control for

Observation

Fine attitude determination / control is performed with FPC-G (other

than Coronagraph mode), and/or C-FPC/C-TTM (Coronagraph mode).

Attitude Maneuver For Pointing direction change or orbit maneuver, satellite attitude

change is performed by RW control.

Orbit Maneuver For maintaining satellite orbit about L2 point and for correcting initial

injection error by launch vehicle, thruster control is performed.

Angular Momentum

Unloading

Angular momentum in RW mainly due to solar pressure accumulation

is unloaded by RCS control.

X-ANT Gimbal Control For X-band antenna communication link, AOCS controls its gimbal.

FDIR Fault Detection, Isolation and Reconfiguration


Date: 5/6/2023　


Pointing Correction Pointing correction by C-FPC (acquisition) or C-TTM (unloading)

Table 6.10 AOCS Main Functions

(3)　AOCS　Performances

SPICA mission pointing requirements are listed in Table 6.11. Each of the SPICA mission

observations is exclusively performed and they are divided into two groups: “with instruments

except for coronagraph” and “with coronagraph”. Pure AOCS realizes the performances listed in

Table 6.12 including alignment error and internal disturbance error. By incorporating FPC-G into

ADS, the performance can be improved as listed in Table 6.13 defined as FPC-G mode. This

performance is applied to the observations except for the coronagraph observation. Figure 6.14 lists

the performance for coronagraph observation that can be realized by C-FPC and C-TTM. In this

mode, AOCS mode performance listed in Table 6.12 is assumed for C-FPC acquisition and C-TTM

unloading.

Observation Mode Absolute Pointing Accuracy Pointing Stability

Except for Coronagraph 0.135 [arcsec] (3σ) 0.075 [arcsec] (0-P) / 200sec

Coronagraph 0.03 [arcsec] (3σ) 0.03 [arcsec] (0-P) / 20min

Table 6.11 Mission Pointing Requirements

Note: 0-P = 0 to peak (half the width of pointing fluctuation over a given period of time)

Error FactorAbsolute Pointing Accuracy

[arcsec] (3σ)

Pointing Stability

[arcsec](0-P) / 20 [min]

Remark

Alignment Error 10 0.00

Determination Error 10 0.45 (*1)Control Error 0.02 0.02 (*2)Internal Disturbance 0.03 0.03

Total (LS) 20.05 0.5

Requirement 30 0.5

Table 6.12 AOCS Mode Pointing Performance

(*1) STT / IRU ADS (*2) AOCS control Note: 0-P = 0 to peak (half the width of pointing

fluctuation over a given period of time)


Date: 5/6/2023　



[arcsec] (3σ)

Pointing Stability

[arcsec](0-P)/200 [sec]

Remark

Alignment Error 0.065 0.005

Determination Error 0.02 0.02 (*3)

Control Error 0.02 0.02 (*4)

Internal Disturbance 0.03 0.03

Total (LS) 0.135 0.075

Requirement 0.135 0.075

Table 6.13FPC-G Mode Pointing Performance

(*3) STT / IRU ADS with FPC-G (*4) AOCS control Note:0-P=0 to peak


[arcsec] (3σ)

Pointing Stability

[arcsec](0-P) / 20 [min]

Remark

Alignment Error 0.005 0.005

Determination Error 0.01 0.01 (*5)

Control Error 0.01 0.01 (*6)

Internal Disturbance 0.005 0.005

Total (LS) 0.03 0.03

Requirement 0.03 0.03

Table 6.14 Coronagraph Mode Pointing Performance

(*5) C-FPC detection (*6) C-TTM control Note:0-P=0 to peak

(4)　AOCS　Mode　Diagram

In order for initial attitude acquisition after the separation from a launch vehicle, mission

observation, orbit maneuver, and FDIR, AOCS has several modes as shown in Figure 6.10.


Date: 5/6/2023　


Fig.6.10 AOCS Mode Diagram

6.8　Reaction　Control　Subsystem

The SPICA reaction control subsystem (RCS) is required not only to control the attitude of the

spacecraft but also to correct the trajectory toward Lagrange point (L2), and to inject the spacecraft

into a Halo orbit around the L2. The total required delta-V is 110m/s. Table 6.15 shows the

requirements for the SPICA RCS.

To comply with such requirements, we adopt a monopropellant propulsion system in the SPICA.

This RCS has 4 23N monopropellant (Hydrazine) thrusters, which enable the trajectory control

maneuvers and orbit injection. The system also requires 8 3N monopropellant (Hydrazine) thrusters

for orbit stabilizing and attitude control maneuvers.


Date: 5/6/2023　


6.8.1　System　Description

(1)　Overview

The RCS consists of 4 propellant tanks, 4 23N thrusters, 8 3N thrusters, piping, valves, filters, and

heaters. Both types of the thrusters work under the blowdown pressure. Heaters and temperature

sensors are installed to preheat the thrusters and provide thermal control of valves, tanks and piping.

The system block diagram is shown in Fig 6.11.

Thrusters are installed in some thruster modules, and latching valves and filters in some valve

modules. All connecting points along the pipe lines from the tanks to the thrusters are completely

welded to preclude any possibilities of leakage of the propellant and pressurant gases.

(2)　Tank　Module

The RCS has tanks with bladders or EPR diaphragms or the like installed inside. The tanks, made

of alloy (Ti-6A1-4V), carry fuel and pressurant gas. The tank shell is formed by superplastic forming

process and fabricated by electron beam welding.

(3)　Thruster　Module

The RCS thrusters can generate 3and 23N class thrust in the blowdown. Each monopropellant

thruster consists of a hydrazine decomposition chamber (catalyst bed), a nozzle, a thruster valve etc.

Main materials are Hanes alloy 25 and SUS 304, and S405 catalyst for the hydrazine decomposition.

A heater and a temperature sensor are installed to every chamber so as to heat the catalyst bed to

about 100 ℃ prior to firing. A space-proven double-seated valve is adopted for the thruster valve.

6.8.2　Performance

The performance specification of the RCS is listed in Table 6.16. Expected dry weight of the RCS is

100 kg, and the weight of initially loaded propellant is 220 kg.

Table. 6.15 Requirements for Propulsion System

Operation Item on Propulsion system Phase Δ V (m/ s) Impulse (Ns) Requirement

Initial attitude acquisition Separation from Rocket 50 10000Torque generation forSun acquisition (Y axis) andEarth acquisition (-Z axis)

Course correction #1 1day after

Course correction #2 10 days after

Attitude control onTransfer orbit Transfer orbit Torque generation for

Sun direction control (Y axis)

Halo orbit injection at L2 Orbit injection 30 0 Δ V for Halo orbit injection

Halo orbit keeping 10 0 Δ V for Halo orbit keeping

Wheel unloading 0 30000 Unloading of external torque forAccumulated at wheel

Attitude keeping atSafe hold mode 0 30000 Change to evacuation attitude at

Failure occurrence

- - Contamination control andExcess translational force

10 30000

Δ V for error modification atInitial course

Through the mission life

Observation (steady)

Table 6.16 Performance of Propulsion System

Operation Item on Propulsion system Requirement Recital Max Impulse 65000 [Ns]/ 1 time Δ V = 25.1 [m/ s] or more Impulse Resolution 25 [Ns] or less Δ V = 0.01 [m/ s] or less Max Torque 8 [Nm] Angular Momentum 0.1 [Nms] = 0.001 [deg/ s] or less Direction of Translational Force +Z (Satellite Coordination) Mission Life 5 years

Propulsion System Mono-propellant(Hydrazine) / Blowdown

EPR Bladder/ Diaphragm

Thruster Configuration 3N thruster× 8 23Nthruster× 4

Specific Impulse 210[s] : 3N thruster 230[s] : 23Nthruster Continuous Combustion

Propellant Weight 220.0 [kg] 10% Margin Propulsion System Weight 100 [kg]

Fig. 6.11 Flow Diagram of Propulsion System

GFDPFD ： Fill / Drain Valve

PRS ： Pressure Transducer

FLT ： Filter

LAV ： Latching Valve

23N ： 23N Thruster

3N-A ： 3N Thruster A-line

3N-B ： 3NThruster B-line

Legend

3N-A 3N-B23N 23N

TP1

TP2

TP3

TP4

TP1

TP2

TP3

TP4

FLT

PRS

PFD

TNK1

TNK2

TNK3

TNK4

GFD1

GFD2

GFD3

GFD4

7.　Payload　Configuration　

7.1　Cryogenic　System

7.1.1　Thermal　Insulation　and　Radiative　Cooling　System　(TIRCS)

The SPICA adopts a new concept of cryogenic system that uses no cryogen. The 4.5 K stage,

consisting of STA, TOB, IOB and some of focal plane instruments (FPI), is refrigerated by the

combined method of mechanical cooling and efficient radiative cooling in the stable thermal

environment at the Sun-Earth L2. In consideration of the cooling capacity of a 4K-class mechanical

cooler (4K-MC) at the end of life (EOL) and on the assumption of Joule heating of the FPI being

approximately 15 mW, the baseline design of the thermal insulation and radiative cooling system

(TIRCS) was determined by the thermal and the structural analyses, so that the parasitic heat flow

from the hot stages to the 4.5 K stage can remain at less than 25 mW(TBD).

The 4.5 K stages such as STA, TOB, IOB and FPI are surrounded by TIRCS consisting of a baffle,

a telescope shell, three shields and a sun shield to reject the heat flow from the outer environment, as

depicted in Fig. 7.1 and 7.2. The Multi-Layer Insulation (MLI) is attached to the sun shield and the

shield #3 to block thermal radiation, while radiator part of each shield remove most of the absorbed

heat from the sun and the Bus Module (BM) to deep space. The layout of these shields and the solar

array paddles is determined to optimize radiative cooling. The STA, the telescope shell and the three

shields are structurally supported by the BM with main trusses between them. The trusses are made

of the carbon fiber reinforced plastics (CFRP) and the alumina fiber reinforced plastics (ALFRP)

with low thermal conductivity, while the baffle and the main part of the telescope shell are made of

high thermal conductivity CFRP to dissipate the heat flow. The sun shield and three shields of

aluminum plates, and structural frames are connected by thrust trusses of ALFRP as well. The shield

#3 and the baffle are supported axially by the BM and the telescope shell, respectively. Wire harness

between FPI and electronics equipments in the SVM is assumed to be made of Manganin and to be

1000 lines. The heat rejection system (HRS) transports exhausted heat from mechanical coolers on

the cooler base plate to the cooler radiator. A loop heat pipe, which has some advantages such as

flexibility in components arrangement and adaptability for long-distance transportation of large heat

amount, is assumed to be a heat transport device.

Thermal analysis was carried out under the analytical conditions listed in Table 7.1. The thermal

radiation between cold stages is much smaller than that between hot stages because of the small

absolute values of the temperatures; the heat flow between cold stages is determined mainly by

thermal conductivity through structural supports and wire harness. A result of heat flow analysis for

the steady state shows that the total amount of parasitic heat and heat dissipation is less than 40

mW(TBD) at the 4.5 K stage, while temperature of the baffle and the telescope shell are lower than

14 K(TBD) and 30 K(TBD), respectively. Average temperatures of the sun shield and three shields

are approximately 158 K, 50 K, 75 K and 110 K, respectively. On the other hand, , the bread board

model (BBM) for an upgraded 4K-MC successfully demonstrated cooling power larger than 50 mW,

which indicates the results of the thermal analysis meets the design requirements. And then, a result

of preliminary time-dependent analysis gives a transient cooling profile with 2 sets of 4K-MC at the

initial phase as shown in Fig. 7.3. It shows that the cooling time of 168 days is required to reach

lower than 5 K after launch.

Bus Module Panel

SAP

Shield #1

Shield #2

Sun ShieldTelescope Shell

Baffle

Radiator

Shield #3

Radiation Shield

Bus Module Panel

SAP

Shield #1

Shield #2

Sun ShieldTelescope Shell

Baffle

Radiator

Shield #3

Radiation Shield

Fig. 7.1 Baseline configuration of a SPICA spacecraft

Table 7.1 Thermal analytical conditions

Parameter Value

Space background 3 K (fixed)

Upper panel of Bus Module (BM) 253 K (fixed)

Solar array paddle (back side) 373 K (fixed)

4.5 K stage (STA, TOB, IOB, FPI) 4.5 K (fixed)

Solar heat flux density 1376 W/m 2

Exhausted heat from mechanical coolers 600 W@293 K

Heat dissipation from FPI 15 mW (fixed)

Wire harness 0.1 mm x 1000 wires

Inner ShieldMiddle ShieldOuter Shield

Fig. 7.2 Schematic drawing of cryogenic system

0.0

50.0

100.0

150.0

200.0

250.0

300.0

0 50 100 150 200 250 300 350)時間（日

K温

度（

）サンシールド

#3シールド

鏡筒ＦＰＩ、主鏡、冷凍機

Temperature (K)

Time (day)

Telescope shell FPI, Primary mirror,

4K -J T cooler

Shield #3

Sun shield

(a) Cooling profile in the overall temperature range

Outer Shield

0.0

5.0

10.0

15.0

20.0

25.0

30.0

0 50 100 150 200 250 300 350)時間（日

K温

度（

）鏡筒

冷凍機

主鏡

FPI

168日

Telescope shell

4K-J T cooler

Primary mirror

Time (day)

Temperature (K)

168 days

(b) Cooling profile in the lower than 30 K range

Fig. 7.3 Cooling profile of the 4.5 K stage

7.1.2　Mechanical　Cooling　System　(MCS)

The SPICA telescope allows high sensitivity and long observation time owing to advanced

mechanical cooling system (MCS). The 4K-MC for the 4.5 K stage is a 4K-class Joule-Thomson

cooler (4K-JT) connected with a precooler of a 20 K-class two-stage Stirling cooler (2ST),which

does not use consumable cryogen. Far-infrared instruments such as SAFARI on the IOB require

further cooling by 1.7K with the 1K-MC consisting of a 1K-JT with 3He working gas and a 2ST

precooler. Specifications of mechanical coolers for the SPICA are listed in Table 7.2.

Based on heritage of 2ST onboard the successful AKARI, 2ST for SPICA is required to have

higher reliability for continuous operation during more than 3 years to complete the mission. At the

same time, the cooling capacity at 20 K has to be increased, because it also greatly contributes to t

increasing cooling capacity of the 4K-JT and the 1K-JT. The engineering model of upgraded 2ST,

which uses lower outgassing materials and less amount of glue, is shown in Fig.7.4. It was proved to

provide higher cooling capacity of 325 mW(TBD) at 20 K. This upgraded 2ST is to be used for the

Soft X-ray Spectrometer (SXS) onboard the next Japanese X-ray astronomy satellite Astro-H

launched in 2013. The cooling performance for the 2ST engineering model (EM) was examined in

various conditions for the Astro-H/SXS, and the long-life test has been also performed to verify the

reliability at EOL for the Astro-H/SXS and the SPICA.

Based on heritage of the 4K-MC with the cooling power of 25 mW, which was developed for the

Superconducting Submillimeter-Wave Limb-Emission Sounder (SMILES) mission at the Japanese

Experiment Module Exposed Facility (JEM-EF) of the International Space Station (ISS) to be

launched in 2009, the upgraded 4K-MC BBM was designed and fabricated by combining new JT

heat exchangers of coaxial double tubes with low pressure loss with the modified 2ST. Test results

with this BBM show that the maximum cooling power of 50.1 mW was efficiently obtained with an

electric input power of AC 55.9 W for the JT compressors and AC 89.2 W for the 2ST. The

remarkable improvement of the cooling power at the 4.5 K stage is attributed to the increase of the

mass flow rate realized by high-power precooling at 18 K by the modified 2ST with the extended 8-

mm-diameter second displacer.

The 1K-MC has been developed for the 1.7 K stage of SPICA. Test results for the 1K-MC BBM

indicate that the cooling power was successfully improved by 16.0 mW with an efficient input power

of AC 76.6 W for JT compressors and AC 89.0 W for the powerful 2ST. Higher mass flow rate,

obtained by cooling the 3He gas in the JT circuit at 12 K by the modified 2ST, drastically increases

the heat lift capacity at 1.7 K. Since this 1K-MC is also to be used for the Astro-H/SXS, the cooling

performance of the 1K-MC EM has been tested in various conditions for the Astro-H/SXS as well as

2ST. The long-life test is scheduled to start to verify the reliability at EOL for the Astro-H/SXS and

the SPICA.

It is notable that all the coolers satisfy the cooling requirements at the beginning of life (BOL),

whereas long life tests for all the coolers are under preparation for reliability verification at EOL. In

the MCS as depicted in Fig. 7.5, two sets of 4K-MC and 1K-MC are employed for redundancy,

respectively.

2ST 4K-MC 1K-MC

Cooler type 2-stage Stirling JT with 2ST 3He-JT with 2ST

Cooling object Exclusive for JT STA, IOB and FPI Far-IRInstrument

Cooling requirement

200mW@20K(EOL)

[email protected](EOL)

[email protected](EOL)

Driving power < 90 W < 160 W < 180W

Heritage AKARI(2006-)

ISS/JEM/SMILES(2009-) N/A

R&D levelEM for

Astro-G (2013) Astro-H (2013)

BBM for SPICA EM for Astro-H

Table 7.2 Specifications of mechanical coolers for SPICA

Fig. 7.4 Upgraded 2ST cold head

Fig. 7.5 Mechanical cooling system

7.2　STA　

The baseline model of the SPICA telescope consists of a Ritchey-Chrètien type telescope with a 3m

class (current base line: 3.5m) single aperture primary mirror, which is the largest size that the rocket

fairing can accommodate (see Fig. 7.6). The telescope system includes the telescope optical bench

(TOB), which is possibly separate from the instrument optical bench (IOB), the focusing mechanism

at the secondary mirror, and the baffling system that prevents stray light from reaching the

instruments and thus allows for an unprecedented sensitivity of SPICA observations in the core

wavelength range (5–210m). Depending on the material of the TOB, a stress-relief mechanism

may also be required for the mirror supports. It is required to be diffraction-limited at 5m and

optimized for the mid- to far-infrared wavelength range. Basic requirements are summarized in Table

7.3, while detailed scientific requirements are described in the SPICA telescope science requirement

document (SciRD). The mirrors will be made of silicon carbide (SiC) or its related material, which

has a large heritage of the AKARI and Herschel telescopes.

Fig. 7.6 Baseline design of the SPICA telescope assembly

Parameter Baseline specification or requirements

Telescope configuration Two-mirror on-axis Richey-Chrètein

Primary mirror size 3m class diameter (current baseline 3.5m)

Operating temperature < 6K

Effective focal length ~18m

Field of view 24’ (diameter)

Image quality Diffraction limited at 5μm at operating temperature

Equivalent to the wave front error (WFE) < 350nm rms within 5’

radius

Surface roughness < 20nm rms

Core spectral range 5–210μm

Mirror reflectivity >97.5% above 30μm and >95% below 30μm

Total mass <700kg

Table　7.3　Baseline　specification　and　basic　requirements　for　the　SPICA　telescope　system

ASF TAS-F

Fig.7.7 ASF and TAS-F

STA is supported by main trusses of Cryogenic Assembly (CRYO) at 8 interface points. Main truss

consists of 3 stages, which is made of CFRP truss and ALFRP truss respectively as shown Fig.7.8.

The influence of main trusses thermal distortion on WFE is one of the interface issues between STA

and CRYO. Preliminary thermo-elastic analysis conducted by ESA shows the design of main truss

meets the WFE requirement.

Fig.7.8 CRYO main truss

7.3　FPI　

7.3.1　Overview　

The Focal Plane Instruments (FPIs) onboard SPICA shall be attached to the STA via the Instrument

Optical Bench (IOB, see Figure 7.9) which is thermally lifted at ~4.5K by a J-T cooler.

The FPIs consists of the following:

MIRACLE (Mid-InfRAred Camera w/o Lens)

MIRMES (Mid-IR medium-resolution echelle spectrometer)

MIRHES (Mid-IR high-resolution echelle spectrometer)

SCI (SPICA Coronagraph Instrument)

SAFARI (SPICA Far-infrared Instrument)

BLISS (Background-Limited Infrared-Submillimeter Spectrograph)

FPC (Focal-plane finding camera) #FPC-G is a guider camera for the attitude control.

*The instruments mentioned above are potential candidates, but still subject to a future selection.

Fig. 7.9 IOB and the FPI volume

Only one FPI (excluding FPC-G) shall be in operation for astronomical observations. FPC-G will

be also in operation for most of the observations (see chapter 4). Cold mass of 150kg (including 20%

margin) is allocated for the overall FPIs. JAXA and ESA agreed to allocate 50kg (including 20%

margin) to SAFARI, the European contribution. Cooling power constraint is also tight : only 15mW

at 4.5K stage, 5mW at 1.7K stage lifted by the J-T coolers. At 4.5K stage, FPC-G (routinely operated

instrument) plus all the parasitic loads amounts to approximately 2mW. Since the resource allocation

for FPI is so stringent, we will make a selection of FPIs before the SDR (in Winter 2009 - 2010). The

selection process and criteria are under discussion. The current cold-mass allocation plan is shown in

Table 7.4. Note that BLISS, a candidate of the US contribution, is excluded from the current mass

allocation.

FPI Weight (kg) comment

MIRACLE

60.0

with 20% margin

MIRMES

MIRHES

SCI 30.0 FPC 10.0 SAFARI 50.0 BLISS 30.0 (optional FPI)

IOB, FPI cover 50.0 with 20% marginTotal 200 excluding BLISS

Table 7.4 FPI Cold mass allocation (excluding BLISS)

IOB is modeled to be a disk made of Ceramic material, 2000mm in diameter and 70mm in

thickness. It is supported by 3 titanium bipods. Analysis shows its first natural frequency higher than

80Hz which is required to be decoupled from satellite first natural frequency.

IOB is still subjected to a more detailed mechanical design.

Fig. 7.10 shows planned field-of-view (FoV) configuration (projected on the sky). Within the

unvignetted field of STA (30 arcmin in diameter), all the FoVs, including two redundant FoVs of

FPC shall be placed. A possible design of the pick-off mirror location is also show in Fig. 7.11.

Fig. 7.10 A planned configuration of the field-of-views (FoVs) of FPIs, projected on the sky. Those

of MIRMES, MIRHES are planned to be placed nearly contiguously to the MIRACLE ones.

BLISS’ FoV is not shown.

Fig. 7.11 A possible design of pick-off mirrors below the IOB.

7.3.2　Mid-Infrared　Camera　:　MIRACLE　

MIRACLE( Mid-InfRAred Camera w/wo LEns ) is a focal plane instrument for wide field imaging

and low-resolution spectroscopic observations (R=/) over a wide spectral range in the

mid-infrared (5-40m). MIRACLE consists of two channels (MIR-S and -L), both of which have

mostly the same optical designs. Each of them has fore-optics which makes images of telescope

focal plane. A field mask wheel is installed at the focal plane in order to provide optimal slits in the

spectroscopic mode. Subsequent camera optics has a pupil position where filter wheels with band-

pass filters and grisms are installed.

A successful optical design with only reflective components (i.e. without lens) is shown in Fig.

7.12. Diffraction-limited performance is achieved at 20 micron covering the entire 6x6 arcmin FOV.

The fore-optics provides the re-imaging focus with diffraction-limited image quality covering the

entire 6x6 arcmin FOV at the shortest wavelength (5 micron). The fore-optics may be shared with

other MIR instruments such as short-wavelength camera and spectrometers.

MIR-S MIR-L

Wavelength coverage 5 m - 26m 20 m - 38m

detector Si:As 1024x1024 Si:Sb 1024x1024

Pixel scale 0.36”/pixel 0.36”/pixel

FOV 6’ x 6’ 6’x6’

Spectral resolution (R=/ ~10 (imaging)/~100 (spec.) ~10 (imaging)/~100 (spec.)

Band pass filters/grism Band pass filters/grism

Size (cm) 125(R) x 50(H) x 30degree 125(R) x 50(H) x 30degree

Table 7.5 Specifications of MIRACLE

Fi

g. 7.12 Field of View of MIRACLE (left) and the schematic diagram of its optical layout

Fig. 7.13 A successful optical design of MIRACLE. The folding mirror for beam pick-off at the STA

focus is omitted for simplicity. The final F-number is set to fast value (3.15) in order to match the

PSF size to small pixel size (18 m) of the detector.

120.0 mm

7.3.3　Mid-IR　Spectrometer:　MIRMES　&　MIRHES

(1) MIRMES (Mid-Infrared Medium-Resolution Echelle Spectrometer)

MIRMES makes medium-resolution (R=/=900~1500) spectroscopic observations over a wide

spectral range in the mid-infrared (10-40m). MIRMES consists of two arms, ARM-S and ARM-L.

They share the same field of view (FOV) area on the focal plane by means of the dichroic beam

splitter. Each arm also has the field mask to limit the FOV and each arm has an image slicer as an

integral field unit (IFU).

ARM-S ARM-L

Wavelength coverage 10.32 m - 19.35m 19.22 m - 36.04m

Spectral resolution (R=/ ~1500 ~900

pixel scale 0.37” 0.72”

slit width 1.11” (3 pixel) 3.6” (5 pixel)

FOV size 12.95” x 5.55” 25.2” x 18.0”

(35pixel x 3 pixel x 5rows) (35pixel x 5 pixel x 5rows)

Table 7.6 Specifications of the Mid-infrared Medium-Resolution Echelle Spectrometer

Fig. 7.14 Echelle formats of MIRMES/Arm-S and Arm-L.

(2) MIRHES (Mid-Infrared Medium-Resolution Echelle Spectrometer)

MIRHES makes high-resolution (R=/

～ 20000 - 30000) spectroscopic

observations in two wavelength ranges, 4-

8m (S-mode) and 12-18m (L-mode).

Immersion gratings enable high spectral

resolution with small mass and volume

resources. Two independent spectrographs

for S-mode and L-mode are under

consideration.

Fig.7.15 Optical layout of MIRHES S-mode

Short(S)-mode Long(L)-mode

Wavelength coverage 4 – 8m 12 – 18m

Spectral resolution (R=/ 30,000 20,000 – 30,000

Slit width 0.72” 1.20”

Slit length 3.5” 6.0”

Dispersion element ZeSe immersion grating KRS5 immersion grating

Cross disperser Reflective reflective

Table 7.7 Specifications of the Mid-Infrared High-Resolution Echelle Spectrometer

MIRHES

S-mode

7.3.4　SPICA　Coronagraph　Instrument:　SCI　

SCI (SPICA Coronagraph Instrument) is a high dynamic-range imager and spectrometer with

coronagraph optics. The primary target of SCI is the direct observation (imaging and spectroscopy)

of Jovian exo-planets in infrared, whilst circum-stellar disks, other type of exo-planets, Active

Galactic Nuclei, and any other compact systems with high contrast can be potential targets.

The specifications of SCI are summarized in a table below. Though diffraction limited image at 5

m wavelength is a specification for the SPCIA telescope, SCI will observe shorter wavelength with

wave front correction by a deformable mirror and a tip-tilt mirror. For a coronagraph method,

binary-shaped pupil mask is a baseline solution because of robustness against telescope pointing

error, achromatic work (except image size effect scaling with wavelength), and simplicity. SCI will

provide coronagrahic imaging mode and spectroscopy mode. On the other hand non-coronagraphic

imaging and spectroscopy is possible because coronagraph mask is removable. Non coronagraph

mode of SCI will work as a general purpose fine-pixel camera and spectrometer. Especially monitor

observation of exo-planet transit is interesting and complementary study of exo-planet to

coronagraphic mode.　

A InSb detector provides higher sensitivity than a Si:As detector in 3.5-5m wavelength,

observable wavelength is extended to the shorter, and simultaneous imaging will be possible with

two detector. Band-pass filters are used for imaging, and transmissive dispersers (e.g., grism) will be

used for spectroscopy.

Thanks to simple pupil shape and active optics, SCI has potential to perform coronagraphic

observation with significantly higher contrast than coronagraph of JWST, which can be a unique

Fig. 7.16 Design of SCI. A tip-tilt mirror and a deformable mirror is used. All device before focal plane mask is made of mirror optics (i.e., no transimissive device). Mechanical changer is used to realize coronagraphic mode and fine camera/spectroscopy mode without mask.

capability to characterize atmospheric feature of exo-planets in 2010s. It should be noticed that SCI

can be useful for monitor observation of exo-planet transit.

Wavelength 3.5-27m (shorter wavelength is not high-contrast

coronagraphic, but sensitive by InSb detector)

Coronagraph method Binary shaped pupil mask

Observation mode Coronagraphic imaging

Coronagraphic spectroscopy

W/O coronagraph imaging

W/O coronagpraph spectroscopy

contrast 10^-6@PSF

Inner working angle 3.3 lambda/D

Outer working angle 16lambda/D *

Detector A 1k x 1K Si:As array, a 1k x 1k InSb array

FoV High-contrast coronagrahpic FoV: 16lambda/D

Non-coronagraphic (high-contrast)image is available for 1’

x 1’

Spectral resolution ~20 and ~200

Table 7.8 Specification of SCI

7.3.5　Focal　Plane　Finding　Camera　　

FPC consists of two components. One is FPI-G that is used for the attitude control system to

stabilize the attitude with an accuracy of 0.05”. FPC-G is a system instrument and will be operated

continuously during the whole SPICA mission. Another is FPC-S that is equipped near infrared

camera for the astronomical observation. FPC-S has a filter wheel with several filters and dispersive

materials. FPC-G and FPC-S are integrated in one package with beam splitter. The specification of

FPC, and optical design are shown below. The mass of FPC is estimated to be < 10 kg.

FPC-G FPC-S

Wavelength coverage 1.6 m (H band) 2 m -- 5 m

detector InSb, 512x412 HgCdTe, 2Kx2K

pixel scale 0.5” 0.18”

slit width 1.11” (3 pixel) 2.16” (3 pixel)

FOV size 4.3’ x 53.5’ 6’ x 6’

heat generation < 1 mW < 2 mW

Table 7.9 Specifications of the Focal Plane Finding Camera

Fig. 7.17 Optical design of FPC (left) & outer shape of FPC (right)

7.3.6　SAFARI　　

SAFARI(SpicA FAR-infrared

Instrument)is an imaging spectrometer

with both spectral and photometric

capabilities covering the ~33-210m

waveband. The key performance and

instrument-design requirements are

indicated in Table 7.9. The baseline

optical configuration of the instrument

is a Mach-Zehnder imaging Fourier

Transform Spectrometer. The main

reasons for selecting this instrument

configuration over other potential

candidates are, (1) the high spectrophotometric mapping speed of the imaging FTS due to the large

FOV and spectral multiplex advantage, (2) the ability to incorporate straightforwardly a photometric

imaging mode which is a key scientific requirement, (3) this configuration lends itself to a high level

of operational flexibility to tailor the spectral resolution of the instrument to the science programme

and (4) the photon noise from the relatively broad band FTS matches the NEP that will be achieved

by using detector technologies available in the timeframe of the SPICA mission.

There are four competing detector technologies under evaluation for use in the instrument;

Transition Edge Superconducting (TES) bolometers, Ge:Ga Photoconductors, Kinetic Inductance

Detectors (KID) and Silicon bolometers. The rationale for not adopting a baseline detector

technology at this stage of the instrument programme is that more time for development is required

Table 7.10 Key Scientific and design requirements of SAFARI

Fig.7.18 Cold 3-D hardware layout of SAFARI

in order to make a fair tradeoff between scientific potential, technical risk and programmatic

considerations. Detailed description of SAFARI will be provided by the SAFARI consortium in a

separate document.

7.3.7　BLISS　　

BLISS (Background-Limited Infrared-Submm Spectrograph) is an extremely sensitive broad-band

far-infrared and submm spectrometer proposed mainly by astronomers in US. Unless the US

commitment to SPICA will be formally announced, BLISS is assumed to be an optional FPI. BLISS

covers the entire far-infrared and submm wavelengths (38-430mm) with 4200 superconductive

bolometer arrays cooled down to 50mK by a dedicated adiabatic demagnetization refrigerator.

Table 7.11 Bliss Specification

Fig. 7.19 Cold hardware 3-D layout of BLISS, consisting five WaFIRS (2-D optics by using concave

8.　AIV/T　Plan　　

8.1　General　

Since it is difficult to simulate the orbit environment for the SPICA on the ground, system-level test

and calibration are significant challenges for us. Among the biggest is verification of the various

payload opto-mechanical budgets assumption under a flight-like thermal and zero-g environment. To

address this, an existing JAXA cryogenic test facility is to be upgraded to allow optical testing of the

telescope assemblies and a suite of instruments at the operating temperature lower than 6 K (target:

<5K) .

For the Bus Module, the integration and test is performed at the JAXA Tsukuba Space Center*1.

Special test for the attitude control is performed to check the vibration environment.

For the Payload Module, Focal-Plane-Instrument Assembly integration and test are performed at

JAXA Sagamihara Campus*2. Cryogenic Assembly test is performed at the JAXA Tsukuba Space

Center. STA is integrated, and standard room-temperature space environmental tests, acoustic and

vibration tests, and optical tests at temperature lower than 100K are performed at ESA.- After STA is

delivered to the JAXA Tsukuba Space Center, final optical verification is performed at the operating

temperature lower than 6 K (target:<5K). Final focus adjustment is made along with the accurate

measurement of the focal position of the STA by driving an adjustment mechanism that will be

installed into the secondary mirror support.

All of the Bus Module and the Payload Module are integrated at the JAXA Tsukuba Space Center.

Final electric test and acoustic test are performed there before its shipment to the JAXA

Tanegashima Space Center*3.

*1:The Sagamihara Campus, with a view of the surrounding Tanzawa mountains, was established in

April 1989 as the core of the former Institute of Space and Astronautical Science (ISAS). It has

research and administration buildings, a research center, and buildings for the development and

testing of experimental equipment for rockets and satellites.

*2:The Tsukuba Space Center (TKSC), located in Tsukuba Science City, opened its doors in 1972.

The TKSC, which sits on a 530,000 square-meter site, with beautiful natural surroundings, is a

consolidated operations facility with world-class equipment and testing facilities.

*3:The Tanegashima Space Center (TNSC) is located along the southeast coast of Tanegashima in

the south of Kyusyu. It is the largest launch facility in Japan (9,700,000 square meters).

It is known as the most beautiful rocket-launch complex in the world.

The center consists of the Osaki Range, the main range for mid to large-size rockets, the Takesaki

Range for small rockets, static firing test facilities, and tracking facilities.

As for SPICA system, MTM (Mechanical Test Model) / TTM (Thermal Test Model) and PFM

(Proto Flight Model) are tentatively planned to be built.

A dynamically representative structure and a thermally representative structure shall be used for

MTM/TTM tests. As TTM tests is to verify mathematical thermal model and passive cooling

capability, not active cooling capability which shall be verified by individual development tests,

thermal similarity of each component with the flight model shall be considered.

As delivery of STA PFM is expected at the 1st quarter of 2016, environment tests for Payload

Module are performed with STA STM and FPI dummies. Just before Payload Module is integrated

to Bus Module, STA STM and FPI dummies are replaced by their flight models, which complete

tests described in 8.2 and 8.3 respectively. Since STA PFM and FPI PFM have limited access after

integration to Bus Module, critical measurements shall be conducted in Payload Module

environment tests. So STA STM is expected to be mechanically and thermally identical with PFM.

Optical performance is not discussed.

Electrical tests and environmental tests are performed as a SPICA system. Interface verification of

FPIs with their electronics shall be performed at this stage. Configuration for this interface check is

TBD. It is not feasible to keep Payload Module at operational temperature <6K (target: <5K) after

this stage, provisions for health check and monitoring of FPI at room temperature until the launch

shall be considered.

8.2　STA　in　Japan　The verification and performance testing of the STA is challenging, especially at cryogenic

temperatures. The Japan-Europe joint telescope working group has worked on defining how and

where each step of the testing process should be carried out. A current plan for a split of

responsibilities between Europe and Japan on the AIV/T of the STA is as follows:

Pre-delivery　(STA　in　Europe):　standard room-temperature space environmental tests and acoustic

and vibration tests are carried out. Optical testing is performed at temperatures lower than 100 K.

Post-delivery　 (STA　in　Japan): 　final optical verification of the STA is executed at temperatures

close to the SPICA operating temperature (nominal: <6K, target: <5K). Final focus adjustment is

made along with the accurate measurement of the focal position of the STA by driving an adjustment

mechanism, which will be installed into the secondary mirror support.

We plan to adopt a horizontal-axis measurement configuration because the vertical-axis

measurement configuration requires a high test chamber (~20m). We are currently modifying the 6-

m radiometer space chamber in JAXA for this configuration test. The chamber has a working length

of 8 m with an optical window in front, and is cooled down lower than 100 K by liquid nitrogen. The

vibration isolation system of the chamber is carefully designed for the purpose of optical

measurements; the optical bench inside the chamber is founded on the seismic slab. In order to

achieve low temperatures as close as possible to the SPICA operational temperature of <6K, an inner

cold shroud enclosing the telescope system is incorporated into the chamber, which is to be further

cooled down by refrigerators. The concept for this modification is based on the X-Ray Calibration

Facility (XRCF) at Marshall Space Flight Center (MSFC), where the mirrors developed for JWST

were evaluated and its flight segment mirrors were tested at 30 K. Our goal is to achieve evaluation

temperatures lower than 10 K throughout the whole system of the STA within the inner cold shroud.

The measurement configuration currently planned for the STA is shown in Fig. 8.1. The issue of

verification of the alignment and wave-front error in a 1-g environment is open and need to be

addressed. A system of g-cancellers during optical verification is being discussed as a possible

approach for the simulation of a zero-g environment. Another approach would be to use FEA to

model the shape of the optical surfaces deformed under 1-g and verify telescope performance under

the predicted deformation, not under the zero-g condition. The validity of this approach would

depend on how accurately we can model the shape of the deformed mirror. The total wave-front

error of the telescope is measured by an optical interferometer through auto-collimation with

reflecting flat mirrors, as in the test of the AKARI telescope. We use the laser-based Twiman-Green

interferometer system in JAXA, which has been proven to have a high tolerance against vibrations in

the 6-m chamber environment. With little prospect for availability of 3m class (current baseline:

3.5m) flat mirrors, the stitching technique is applied to the 3m class telescope pupil with a sub-pupil

array consisting of two discrete 1-m flat reflectors of glass mirrors. A study of stitching algorithms

based on the least-squares method is currently underway by using numerical simulations. Our

current plan is to place a set of two flat mirrors at the radial positions of 770 mm and 1400 mm from

the center of the mirrors, respectively, and rotate them together along the optical axis by a step angle

of 22.5 degrees to cover the whole aperture of the STA. The STA is fixed in the chamber. An

additional unwanted tilt of the flat mirrors that could occur due to the rotation is measured with auto-

collimators from outside the chamber. The tilt is to be taken into account in the stitching analyses or

corrected by tilting the flat mirrors at the corresponding angles if necessary. It is not necessarily 10

K at the flat mirrors and the movement mechanisms; 100 K might still be acceptable, which makes

the thermo-mechanical design of these parts much easier. The back-focal length of the STA is too

short for the interferometer, and we cannot measure WFE by the interferometer placed outside the

chamber; hence we put it with a 5-axis optical adjustment stage inside the chamber by using a

pressure vessel for it and thermally insulating the system with multi-layer insulators.

Fig.8.1 Schematic view of the cryogenic optical testing of the STA by utilizing the JAXA 6m liquid-

nitrogen space chamber at ~10 K with modifications

8.3　FPI　

Integration and verification plan of the FPIs critically depend on the policy of risk management of

the SPICA mission and the timing of the delivery of the whole FPIs. As for the delivery of the whole

FPIs, the timing should be well before the STA delivery from ESA, since the FPI evaluation test will

be undertaken by using the large cold chamber at Tsukuba Space centre of JAXA, which may also be

used for the STA evaluation at cold. As for the risk management policy, here we tentatively assume

the following:

+ For each FPI, not only a flight model (FM) but also a prototype model (PM ) (a back-up of FM,

having fundamentally the same functions as those of FM) shall be constructed. EMs

(engineering models) are also developed for specific tests as well as I/F check before the

development of FM/PM.

+ If serious problem happens in a FM after the delivery of FPI, its PM will be used in place of the

FM with problems so far as the replacement does not give serious delay in project schedule.

Pre-delivery: function, performances of each FPI shall be evaluated independently. This evaluation /

tests of individual FPI should include a) test at ambient temperature/pressure (e.g. mechanical

environmental test, a part of electronics’ function test, etc.) and b) test at cold temperature (planned

temperature of STA/IOB: 4.5K). Moreover, FPI should survive abrupt evacuation (de-pressure)

suffering at the “warm”(STA and FPI are at ambient temperature/pressure) launch. Basically EM and

FM/PM should be tested in the same scheme and conditions.

For the individual tests at cold temperature, dedicated cold chamber and test equipments should be

prepared. For example, Nagoya University will develop a cold test chamber for MIRACLE, which

will be equipped with cold working volume of 1.5m x 0.5m x 0.5m (enough to install either short- or

long-wavelength channel of MIRACLE) at temperature of 4.5K. The test chamber may also be

equipped with a telescope simulator. This chamber may be also used for evaluation of mid-infrared

spectrometers. As for SCI, ISAS of JAXA will develop a dedicated cold chamber.

Post-delivery: before the delivery of the FM STA, FM (neither PM nor EM) of the whole FPIs are

assembled and tested as described below. PM (or EM) may be also integrated for the I/F check

(TBD)

+ Assembly / Integration: FPI assembly (all FPIs and IOB) is delivered to Tsukuba Space Centre

(TBD) and assembled together. Alignment, as well as the weight, is measured.

+ Verification and Test: firstly at ambient temperature, the FPI assembly is tested at the mechanical

environment test facility at Tsukuba Space Centre(TBD). Then the FPI assembly is installed to the

6 m-diameter radiometer space chamber facility at Tsukuba Space Center (TBD) and the evaluation

at cold temperature will be started. The main purposes of this FPI cold test are:

1) all functional test with flight electronics

2) performance evaluation (noise of infrared sensors, crude optical performance test with

light source )

3) focal check between each FPI and the FPC-G with a concise telescope simulator

4) evaluation of interference between FPIs, especially between FPC-G (always operated) and

one of other FPIs. There would exist potential interferences between the FPIs in the

standby mode and the FPI in operating mode.

After the FPI cold test and the STA cold test (section 8.2), the STA and FPI assembly is integrated

and delivered to the mission integration. During long period before the launch, the maintenance

strategy of as well as the health check of the FPI assembly should be also determined.

9.　Project　Plan　

The SPICA is the JAXA-led mission under international corporation and JAXA is responsible for

all the project. If SPICA is selected as one of the Cosmic Vision programs, ESA will be the prime

project partner and responsible for development of the SPICA Telescope Assembly and SAFARI.

Fig.9.1 shows the international corporation scheme and role of each country.

ESAI/F Management

SAFARI ConsortiumSystem Integration

EuropeanTeams

JapaneseTeams

FPI : SAFARI

ESAManufacturing（test @ 80K）

JAXAIntegration（test @<10K）

JAXASubsystemIntegrator

Japanese Group

Korean Team (TBD)

NAOJ (TBD)

FPI:MIRs＋SCI

JAXA SPICA team: System Integration

SPICA Steering Committee

FPI:FPCFPI:BLISS

NASATeam (TBD)

※ FPI : Focal Plane InstrumentSAFARI : SPICA Far-Infrared InstrumentMIRs : Mid Infrared Insturuments (MIRACLE,MIRMES,MIRHES)BLISS : Background-Limited Infrared-Submikimeter SpectrographFPC : Focal Plane finding Camera

Science Advisory Committee

Fig.9.1 International Corporation Scheme

SPICA Steering Group is established early in the program prior to the formal agreement by the both

space agencies. Fig.9.2 shows the SPICA Management Structure. This Steering Group, which is

chaired by a JAXA representative, consists of members from ESA, JAXA and potentially other

national ESI funding agencies. During the Project Formation Phase, the content of the formal

agreement between the two agencies is drafted by the Steering Group (i.e. agreement on the details

of the ESA deliverables and overall joint SPICA program level agreements). The second role of the

Steering Group, which remains in place until the end of SPICA operations, is to monitor the overall

technical, financial and programmatic status of the mission. It has the authority to make high level

tradeoff decisions when required (cost, schedule, risk, science performance etc.) and take executive

actions to deal with problems arising throughout the program. Also Science Advisory Committee is

established. This Advisory Committee, which is chaired by The SPICA Project Scientist, consists of

members from scientists representing the important fields of astronomy. The role of the Advisory

Committee is to review and give advice to SPICA project from scientific points of view.

Fig.9.2 SPICA Management Structure

JAXA SPICA pre-project team is working with three industries at Phase-A: Sumitomo Heavy

Industry (SHI), NEC, and MELCO. SHI is the only industry which has advanced space Cryo

technology. Two other industries are the satellite system experts, each of whom is working with

JAXA individually. JAXA has contractual relationships with them to obtain cooperation in the R&D

from different approaches. JAXA will define the system requirement before SRR, using industries

report. After SRR, JAXA will submit the Request for Proposal (RFP) to industries and select a prime

contractor.

Fig.9.3 shows the Definitions of SPICA Lifecycle and Reviews.

Fig.9.4 shows the Document Tree. SPICA Mission Requirement Document (MRD) is the top level

document and this document defines all the requirements from other requirement documents such as

SPICA Telescope Requirement Document and STA Science Requirement Document. SPICA Project

plan and SPICA Systems Engineering Management Plan (SEMP) is also the top level document.

SEMP is a part of SPICA Project Plan and defines the scope of project and engineering process.

Below the SPICA Project Plan, there are 5 plans. (SPICA Master Schedule, SPICA Tracking and

control Plan, SPICA Science Plan, SPICA System Safety Program Plan, and SPICA Risk

Management Plan.) Interface Control Specification (ICS) Documents is used for JAXA-ESA

interface discussion. ICS-STA is for SPICA Telescope Assembly and ICS-FPI for the Focal Plane

Instrument. Each ICS has two parts and Part-A for JAXA to specify the interface and Part-B for ESA

in reply to Part-A. SPICA Standards are basically from JAXA's standard, but some from ECSS as

required from ESA.

　　　　　 SPICA MDD　 Reference: JAXA_SPICA_SYS_MDD_001

Date: 5/6/2023Issue: 1-0 Page:　82

Phase-B Phase-C Phase-F

概念設計(Concept Development)

計画決定(Project Formation)

基本設計(Preliminary Design)

詳細設計(Final Design)

製作・試験(Production and Test)

射場整備(Launch operations)

初期運用(Initial Mission Operations)

定常運用(Nominal Mission

Operations)後期運用

(Extended Mission Operations)ミッション終了(Mission End)

Phase-EPhase-A

宇宙開発委員会(Space Activity Committee)

理学委員会(Space Science Steering

Committee)

J AXA経営審査(J AXA Management Review)

J AXA技術審査(J AXA Technical Review)

ESA審査(ESA Review)

ライフサイクル(Life Cycle)

Phase-D

SPICAタスクフォース(SPICA Task Force)

プロジェクト準備審査(Project Readiness Review )H20.5.12

プロジェクト移行審査(Project Approval Review)

Cosmic VisionDown Selection

SDR)システム定義審査（PDR)基本設計審査（ CDR)詳細設計審査（開発完了審査打上げ準備完了審査

LRR)（定常運用移行審査定常運用終了審査ミッション終了審査

Cosmic VisionFinal Selection

Preliminary Design ReviewPDR)（

Critical Design ReviewCDR)（ System Acceptance Review

開発研究移行事前評価開発移行事前評価事後評価

SRR)システム要求審査（ FMDR)ミッション最終定義審査（

研究開発研究開発運用

Fig.9.3 Definitions of SPICA Lifecycle and Reviews


Date: 5/6/2023　


J AXASPICA文書

Level 3

J AXASPICA文書

Level 2

J AXASPICA文書

Level 1SPICA衛星プロジェクト計画書

(Project Plan)

SPICASystems Engineering Management Plan（SPICA SEMP）

SPICAシステム安全

プログラム計画書(System Safety Program Plan)

SPICAMission Requirement

Document（MRD)

SPICA総合システム運用コンセプト

SPICA研究コンセプト

SPICA Telescope Requirement

DocumentRSSD-SPICA-00

SPICA総合システム要求書

SPICA衛星システム仕様書

SPICA研究システム仕様書

SPICA研究計画書

(Science Plan)

SPICA追跡管制計画書

(Tracking and Control Plan)

SPICAリスク管理計画書

(Risk Management Plan)

SPICAマネージメントリスク識別情報

SPICA衛星システムリスク識別情報

SPICA追跡管制システムリスク識別情報

SPICA研究システムリスク識別情報

SPICA追跡管制システム仕様書

Interface Control Specification


(ICS-STA Part A)

SPICA設計基準書



(SPICA Standards)

J AXA設計標準

ESA設計標準ESA設計標準ESA設計標準

(ECSS)

SPICA総合システム検証計画書

(SPICA Verification Plan)

SPICA衛星システム組立・試験計画書

SPICA追跡管制システム試験計画書

SPICA研究システム解析計画書

プロジェクトマネジメント規定

J AXA文書 ESA文書

品質マネジメント規定

SPICAマスタスケジュール

(Project Master Schedule)

J AXA設計標準J AXA設計標準

J AXASPICA文書

Level 4

J AXA総合事業計画書

プロジェクトマネジメント実施要領

システム安全標準リスクマネジメント

ハンドブック

周波数管理規定

安全データパッケージ

安全審査実施要領

教育・訓練計画書

教育・訓練実施/ 参加者記録

監査計画書

監査報告書

報告書(問題事項）

事故報告書、ヒヤリ・ハット報告書

(必要な場合)

周波数管理要領

ICSに関する根拠資料・参考資料等

（ICS Reference Documents）

■J AXA- SPICA- IF001 Rev.2.1SPICA INTERFACE CONDITION FOR SPICA TELESCOPE ASSEMBLY (STA)■J AXA- SPICA- IF002 Rev.1SPICA ENVIROMENTAL CONDITION FOR SPICA TELESCOPE ASSEMBLY (STA)■J AXA- SPICA- CRYO0002SPICA OVERVIEW DRAWING OF CRYOGENICS SUBSYSTEM (CRYO)■J AXA- SPICA- MISSION0002SPICA MISSION OVERVIEW

Level1：ベースライン文書Level2：計画書、I/ F仕様書 ※ Level1と2の文書はコンフィギュレーション管理対象

Level3：詳細な仕様書、計画書、基準書等Lebel4：その他 ※ Level4までの文書は主に審査等でSPICA関係者外へ開示されるものであり、プロジェクト関係者のみで扱う文書はこの体系に記述されない

STA Science Requirement

DocumentRSSD-SPICA-00

Interface Control Specification Focal

Plane Instrument (FPI)(ICS-FPI)

Interface Control Specification


(ICS-STA Part B)

Interface Control Specification SAFARI

(ICS-SAFARI)

Interface Control Specification BLISS

(ICS-BLISS)

Interface Control Specification FPC

(ICS-FPC)

米国文書

韓国文書

Fig.9.4 SPICA Document Tree