Solar B - EIS
MULLARD SPACE SCIENCE
LABORATORY
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UNIVERSITY COLLEGE LONDON
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Author: A P Dibbens
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SOLAR B - EIS ICD DOCUMENT
Document Number: MSSL/SLB-EIS/SP003.04 5 July
2000
Distribution:
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NRL
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G Doschek
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C Korendyke
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S Myers
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C Brown
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K Dere
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J Mariska
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NAOJ
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H Hara
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T Watanabe
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RAL
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J Lang
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B Kent
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BU
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C Castelli
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S Mahmoud
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Mullard Space Science Laboratory
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J L Culhane
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A Smith
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A James
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L Harra
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A McCalden
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C McFee
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R Chaudery
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P Thomas
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R Card
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J Tandy
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W Oliver
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P Coker
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R Gowen
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K Al Janabi
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M Whillock
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SLB-EIS Project Office
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A Dibbens
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Orig
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Author:
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Date:
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Authorised By
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Date:
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Distributed:
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Date:
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CHANGE RECORD
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ISSUE
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DATE
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PAGES CHANGED
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COMMENTS
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01
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29 February 2000
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All new
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02
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17 April 2000
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All
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Major update following the engineering meeting in Japan, 6-9 March
2000.
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03
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16 June 2000
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3,4,6,7,8
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Par 4.3.2 units of CLA added. Paras 8.3 & 8.4 added. Par 5.3 updated
to reflect the larger ICU base area in contact with the S/C bus. Par 5.4 added.
Co-planarity added in par 4.3.2. Par 4, drawing references to structure and
templates updated; also mass, M of I, c of g and stiffness properties updated.
Par 9, Power Budget updated. References to cables and connectors added to par
8.1 and Appendix 7 added. Par 4.3.3 changed to reflect 20mm dia shear
pins.
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04
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05 July 2000
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All
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Major revision in preparation for the EIS UK PDR.
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Contents
APPENDICES 1. Structure
GA 2. Mounting Template 3. ICU
Interface 4. Electrical Block Diagram 5. Grounding
Scheme 6. Power Distribution 7. Cables and
Connectors
1 INTRODUCTION
Solar-B will study the connections between fine magnetic field elements in
the photosphere and the structure and dynamics of the entire solar atmosphere.
The mission will perform three basic types of observation with high spatial,
spectral and temporal resolution :
Determination of the photospheric
magnetic vector and velocity fields.
Observation of the properties of the
resulting plasma structures in the transition region and
corona.
Measurement of the detailed density, temperature and velocity of these
structures.
The EUV imaging spectrometer (EIS) will obtain plasma velocities
to an accuracy of <= 10 km s
-1 along with temperatures and
densities in the transition region and corona at <2 arc sec resolution.
2 OVERVIEW
2.1 Description
EIS consists of a multi-layer coated single mirror telescope, and a
stigmatic imaging spectrometer incorporating a multilayer coated diffraction
grating. The image produced by the primary mirror is imaged onto an entrance
slit/slot and the light which passes through this spectrometer aperture is
dispersed and re-imaged at the focal plane of the CCD detectors.
A
separate electronics box (ICU) provides the instrument control functions and
interface with the spacecraft.
For details of the system definition see
RD 3
2.2 System Hierarchy
EUV Imaging Spectrometer
Structure
Mirror Assembly
Grating
Assembly
Slit-Slot Assembly
Clamshell Assembly
Camera
Assembly
Sensors and Heaters
Mechanism and Heater Control
Unit
Thermal Blanket
Instrument Control Unit
Processor
Camera
Buffer
Camera Mechanism Controller
Power Conditioner
Harness
2.3 Block Diagram of
EIS
3 APPLICABLE DOCUMENTS
RD 1 NAO/SLB-EIS/SP/MDP001 MDP-EIS-ICU Electrical Interface
RD
2 MSSL/SLB-EIS/SP/004 Mass Budget
RD 3 MSSL/SLB-EIS/SP/011 EIS System
Definition
RD 4 MSSL/SLB-EIS/PA/003 Cleanliness Control Plan
RD
5 MSSL/SLB-EIS/PA/002 PA Plan
RD 6 SLB-124 Environmental Conditions for
Solar B
RD 7 SR 8189
RD 8 Solar B Electrical Design Standards
(Japan)
4 FILE REFERENCES
The following files are available at the ftp site
indicated
FR1 SOLARB-8193.dxf EIS Spectrometer GA
drawing Birmingham
FR2 SR8154-B.dxf Mounting Template
Drawing Birmingham
FR3 SR8224.dxf Interface drawing for Launch
Lock Birmingham
FR4 Provisional ICU Interface MSSL
Site
addresses:
Birmingham ftp://cad8.sr.bham.ac.uk/pub/solarb/mech
MSSL TBA
NRL TBA
5 SPACECRAFT RESOURCE SUMMARY
The following sections provide the high level status of the mass and
power budgets. No contingencies are held within the EIS project but rather they
are held by the ISAS team and are available through a process of justifiable
request.
5.1 Mass
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Subsystem
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Mass (kg)
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Spectrometer
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57.43
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ICU
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6.0
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Harness
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4.0
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Total
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67.43
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The Instrument Mass Budget is shown in RD 2.
5.2 Power
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Mode
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ICU
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MHC
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CAM
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Mechs Pwr
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Av Power
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Pk Power
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Off
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Off
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Off
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Off
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Off
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0.0
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0.0
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Boot
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On
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Off
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Off
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Off
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14.2
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17.1
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Standby
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On
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Off
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Off
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Off
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14.2
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17.1
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Emergency Safe
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On
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Off
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Off
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Off
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14.2
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17.1
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Manual
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On
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On
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On
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On
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39.8
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55.3
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Auto
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On
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On
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On
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On
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39.8
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55.3
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Engineering
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On
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On
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On
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Off/On
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39.8
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55.3
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Bake-out
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On
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Off
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Off
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Off
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44.2
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44.2
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Note 1. All values are in Watts and refer to primary power.
Note 2. When
the operational heaters are on the power is pulse width modulated. This excludes
the CCD heater which is listed separately
Note 3. The design is such that
operational heaters will be switched off while mechanisms are moved.
Note 4.
Survival power is not included.
Note 5 – The CCD heater will be used to
decontaminate the CCD and will be used with other power systems switched
down
6 SPECTROMETER
6.1 Structure
The subsystems of the spectrometer are supported by a composite structure.
This structure consists of a single base plate which performs the function of an
optical bench. The optical elements are mounted directly (or near directly) from
inserts within this composite base. The Spectrometer enclosure is formed by side
and top composite panels which are held together with titanium inserts. The
upper panel is divided into two parts, one of which is removable to provide
access to the grating and slit-slot assemblies. Access to the mirror is via the
associated end panel. The structure also provides for optical baffling.
The mechanical structure of EIS is shown in the drawing GA Proposal,
SR8193, see Appendix 1 (file reference FR
1).
6.2 Stiffness
The lowest characteristic frequency is 75Hz about the mounting
legs.
6.3 Mechanical
interface
6.3.1 Mechanical details
Details of the mechanical interface of EIS with the spacecraft are shown in
the drawing GA Proposal, SR8193, see Appendix 1 (file reference FR
1).
6.3.2 Specification of attachment surfaces
Attachment surface is titanium insert within a molded carbon fibre
composite
Surface roughness = 1.6
μm CLA
(Centre-Line-Average)
Co-planarity = ±
0.05mm
6.3.3 Attachment Fastening
There are three holes with 3 x 5/16” Unified tapped thread holes with
2xConcentric dia. 16.0 H7 (B0 & C0) and 1xConcentric dia.20.0 H7 (A0) holes
for special shear
bushes. The bolts are 3 x 5/16” Unified tapped
thread with shear bushes 32mm
long, concentrically positioned.
Fastener
Torque: TBD
The required drawing are:
For A0 hole: drawing
SR8217
For C0 hole: drawing SR8218
For B0 hole: drawing
SR8219
6.3.4 Template
The details of the interface template are shown in the drawing Mounting
Template, SR 8154, see Appendix 2 (file reference FR
2).
6.3.5 Launch Lock
While the present baseline does not include a Launch Lock it has been
deemed prudent to provide an appropriate interface if one is introduced later.
This interface is shown in drawing SR 8224 (see Appendix
?).
6.4 Mass properties
The spectrometer mass is provided in section
5.1.
6.4.1 Centre of Gravity
The centre of gravity is at:
x = -0.254m
y = 0.112m
z =
1.69m.
The coordinate system is defined from a local origin in
EIS.
6.4.2 Moments of Inertia
Ixx = 2.29 kg.m²
Iyy = 54.3 kg.m²
Izz = 55.3
kg.m²
6.5 Motors
Electric motors are used in the EIS instrument to move mechanisms. The
following table summarizes their
characteristics:
Table 6.5
Mechanism Characteristics
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Mechanism Subassembly
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Translation
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Actuator
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Encoder
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Average Duty Cycle
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Peak Internal Power
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Average Power
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MIR
Primary Mirror Subassembly
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Coarse Position
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Size 16, 4 phase stepper motor
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Resolver
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2 (20 sec) operations per day
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20 W
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0.0092 W
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Fine Position
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Piezoelectric Transducer
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Strain gauge
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0.5 V step per five seconds
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0.29 W
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<0.05 W
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SLA
Slit/Slot Subassembly
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Slit/Slot Exchange
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Size 12, 4 phase stepper motor
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Resolver
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2 operations per hour
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6 W
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0.0084 W
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Shutter
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Brushless DC motor
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Optical encoder
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1 operation every 5 seconds
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2.65 W
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0.0122 W
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GRA
Grating Subassembly
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Focus Mechanism
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Size 16, 4 phase stepper motor
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Optical encoder
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2 (20 sec) operations per month
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20 W
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0.0092 W
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NOTE: Duty cycle, peak internal power, and average dissipated power values
are preliminary estimates.
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6.6 Field of View and Exclusion Zone
The EIS instrument views the Sun through a front aperture at the end of a
rectangular baffle tube. This baffle tube extends sunward beyond the thin
aluminum filters. The angular size of the Sun is 0.5 degree, and the baffles and
aperture openings are sized to accommodate this angle plus a 2 mm margin all
around.
Figure
6-6a. EIS Entrance ApertureThe front aperture is actually an oval to accommodate the
±10 mm X translation of the primary mirror,
but is assumed to be circular here for simplicity and to be 200.2 mm in
diameter.
While the Sun only occupies a
0.5
° cone angle, the front portion of the
baffle tube serves to protect the thin aluminum filters from micrometeorites,
orbital debris, and contamination. The most likely source of contamination or
damage is from components of the Solar-B spacecraft itself. Outgassing from warm
surfaces and particulates such as paint flakes will be very damaging to the
filters. For this reason, the front baffle tube has been designed so that the
filters have no direct line of sight to other components of the spacecraft. In
the present design, there is a light baffle midway between the filter and the
entrance aperture. A zone of exclusion in front of the EIS aperture is
configured in front of EIS such that no straight-line path within this zone can
reach beyond the middle baffle. Such an exclusion zone has a full angular extent
of 48
° about the S/C Z direction. The entrance
aperture center point location in S/C coordinates and footpoint B0 are given as
S/C coordinates in
Table 6-6. The
coordinates of B0 are as supplied by the System (see file
fairing.pdf
from H. Hara dated 1/18/00), and the EIS aperture has been calculated from this
point, should B0 move, the EIS aperture will move with it.
Figure
6-6b. EIS Baffle Tube and Exclusion
ZoneTable
6-6. Location of EIS Entrance Aperture
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S/C
Coordinate
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Footpoint B0
(mm)
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Center of EIS Aperture
(mm)
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X
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0.0
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104.9
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Y
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605.0
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721.0
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Z
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2767.8
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3161.5
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6.7 Disturbances
Disturbances to the spacecraft can be caused by both translating and rotational
mechanisms.
Translating Mechanisms
- Primary Mirror (MIR): The primary mirror translates back and forth in
the X direction by ± 8mm. This coarsely
positions the image of the region of interest on the spectrometer slit. The
selection of a single mirror telescope forces us to move the 2.3kg mirror
assembly instead of a much smaller secondary mirror to do this. It is driven by
one of the stepper motors mentioned above through a gearhead, which turns a ball
screw. At its top speed of 200 steps/sec, the mirror can move 16 mm in 28.63
sec. It has already been determined that this is UA, but a speed reduction to
100 steps/sec brings us to a CA angular momentum, and further reductions can be
made as necessary. Cell F7 of the spreadsheet contains the time required for a
16 mm translation. The 16mm range is worst case, and corresponds to a slew from
the East limb of the sun to the West limb. A mirror translation will be required
whenever a new target is selected for EIS. This might occur on the order of once
a day.
The cumulative angular momentum for the mirror translation
is independent of the speed of the motion, depending only on the mass of the
object, the distance from the S/C CG, and the distance traveled. For the EIS
Primary mirror, the full range travel results in two UA values. These come into
the CA range when moving smaller distances (on the order of 3mm), but we expect
that since the motion is slow (taking ~1 min or more) the ACS can keep up with
the disturbance. It is expected that re-pointing EIS will only occur when a new
target is chosen for all Solar B instruments and the move can be made while fine
pointing control is not required. If necessary, EIS can delay its move until the
other instruments are finished exposing. It may be necessary to provide advance
information to the ACS system to permit it to anticipate the disturbance. The
S/C ACS engineers in Japan are studying this.
It is important to note that
the range of this motion is limited to ±8mm in
the X direction, so the cumulative angular momentum cannot grow beyond the
maximum reported value. Any movement to an extreme position must be followed by
a movement in the opposite direction. In the long term, the EIS targets will be
randomly located on the sun and the cumulative angular momentum from the primary
mirror will be zero.
- Grating Focus (GRA): A grating focus mechanism is included, and
incorporates the same motor and lead screw combination that is used for the
primary mirror. Focus adjustments move the grating in the X-Z plane along a line
making a 4.3° angle with the Z-axis. The
disturbance was calculated for a 1mm movement of the grating, but normal
movements are expected to be on the order of 0.25mm or less. The movement was
assumed to take 60 sec. It is expected that several focus movements will be made
during the commissioning of EIS, and once a best focus is attained, this
mechanism will rarely be used. The speed of this movement could easily be
increased, as the angular momentum is
small.
Rotating Components:
There are three components that have rotational motions within EIS. The
pivoting of the primary mirror during fine scan mode and the rotation of the
shutter blade both are operational modes that involve continuous periodic
motion. The third, the slit/slot interchange is an intermittent
motion.
- Mirror Fine Scan (MIR): The primary mirror fine scan is a rotation of
the primary by up to 4 arc min about an axis located at the vertex of the mirror
and oriented in the S/C Y direction. During the fine scan, the mirror moves very
slowly, for example in steps of ~1 arcsec/sec. At the end of a scan, the mirror
flies back to the starting position in about 2 sec. This “flyback”
disturbance is calculated here (worst case). However, the sum of the fine scan
and the flyback result is zero cumulative angular momentum for each scan cycle.
The CG of the mirror is 3.81cm from the axis of rotation, so this is treated as
a “static imbalance” disturbance. The disturbances are small since
the angular travel is extremely small.
- Shutter (SLA): The shutter blade is directly driven by a Kolmorgen
1” size stepper motor. The shutter is a thin disk with a sector cut out
for the open position, so we have a static imbalance. The missing mass of this
cutout is the source of this imbalance. Q for this item is 3.89E-6 Kg-m, and the
worst cast rotational speed (15.7 rad/sec) is assumed for a
180° rotation in 0.2 sec in the shortest
exposures. The rotation axis of the shutter is in the X-Z plane and makes a
4.3° angle with the Z-axis. Reversing the
shutter direction periodically possibly after every exposure can null the
cumulative angular momentum.
EIS’s shutter executes one
complete cycle for each exposure, and the cadence can be as high as 1 Hz. This
may be close to a vibration mode of the solar panels. The solar panel mode
frequencies are TBD and must be determined on orbit. Exposure cadences must be
chosen after launch to avoid exciting these resonances.
- Slit/Slot Mechanism (SLA): The slit/slot mechanism is used to bring
different sized apertures to the slit position. There are four apertures located
90° apart on a paddle wheel-like holder. A
90° rotation brings another slit into position.
The mechanism is driven by a size 12 (0.75”) CDA Astro stepper motor. A
reducing gearhead is used to give the positioning accuracy required. The
rotation axis is in the X-Z plane, nearly parallel to X. 14 sec are required to
make the 90° movement, and intermittent
operations several times per hour might occur in some observing plans. The
device is tiny and moves slowly putting its disturbance into the “A”
category.
Conclusions
All the disturbance torques except those associated with retargeting EIS by
moving the Primary Mirror were found to be Acceptable (A) or Conditionally
acceptable (CA). Steps that will be taken to minimize the UA effects on
cumulative angular momentum are as follows: 1) The ACS system will study
compensations for the EIS Primary Mirror movements. 2) Primary movements will be
scheduled, not autonomous, and will occur within periods where fine pointing is
not required by other instruments. 3) Movements of the Primary Mirror will be of
duration consistent with the time constants of the ACS control system. 4) EIS
will keep the moving mass to the practical minimum consistent with mechanical
and optical constraints
6.8 Provisional EIS Co-alignment
The Solar B instrument complement must be sufficiently well co-aligned
to have reasonable overlap of the three instrument’s fields of view. To
accommodate co-alignment at the spacecraft level, the EIS instrument will have
an optical alignment cube. The alignment cube will be aligned to the EIS
telescope optical axis to within 20 arc-seconds on an optical bench in the
laboratory. The cube will be utilized to co-align the EIS optical axis with the
axes of the spacecraft sun sensor and the other instruments. Along the
north-south direction, the EIS field of view is 1024 arc-seconds and along the
east-west direction the coarse mirror motion corrects for up to +/- 800
arc-second. Even with an additive 1 arc-minute tilt of the primary and grating,
the EIS field still maintains complete coverage of the SOT field of view and is
still well centered within the XRT field of view.
Table 10. Preliminary
internal EIS co-alignment error budget (significant contributors only)
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Subassembly
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Significant error budget contributor
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Alignment cube
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10 arc-seconds transfer error to cube front face
20 arc-seconds
mechanical variation
total: 30 arc-seconds optical axis error
note: cube
face tolerance of <5 arc-seconds expected
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Parabolic mirror
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30 arc-seconds optic/mount tilt
30 arc-seconds structure tilt
total:
60 arc-seconds of tilt, 120 arc-seconds equivalent solar error
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Grating
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30 arc-seconds optic/mount tilt
30 arc-seconds structure tilt
total:
60 arc-seconds of tilt, 62 arc-seconds equivalent solar error
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RSS total:
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138 arc-seconds
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Note: Additive error budgeting at the component interface level has been
utilized to simulate a cross-coupled error. The gross errors in each component
were root sum squared to produce the total.
6.9 Thermal interface
6.9.1 Attached Area
Area of attachment points are:
A0: 7.1 cm
2B0: 8.2
cm
2C0: 8.2 cm
2Power flow across attachment point
-5 to +5W
6.9.2 Heat Dissipation Across Attachment
Points
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Off
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Typical operating
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Peak
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Survival
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Absolute W
A0
B0
C0
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Fill in etc.
<2.5
<2.0
<2.0
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<2.5
<2.0
<2.0
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2.5
2.0
2.0
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2.5
2.0
2.0
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Density W/cm2
A0
B0
C0
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<0.4
<0.3
<0.3
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<0.4
<0.3
<0.3
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0.4
0.3
0.3
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0.4
0.3
0.3
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Note that this is required to be < 5W per attachment point in either
direction
6.9.3 Heat Capacity
The heat capacity of the Spectrometer is 40900
W/K.
6.9.4 Acceptable Temperature Ranges
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Critical Item
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Operating
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Survival
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Mirror
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+10 to +30oC
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0 to +40oC
|
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Piezo actuator
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+10 to +30oC
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0 to +40oC
|
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Grating
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+10 to +30oC
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0 to +40oC
|
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CCD
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-55 to +30oC
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-100 to + 60 oC
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Clamshell filter
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TBC
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TBC
|
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Secondary filter
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TBC
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TBC
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Slit-Slot
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+10 to +30oC
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0 to +40oC
|
6.9.5 Positions of Temperature Measurement
The following sensors are provided by the spacecraft at the locations
indicated.
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Sensor/Position
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X
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Y
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Z
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Primary Mirror support
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105.0
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814.38
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204.05
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E-box doubler plate
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-261.0
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668.25
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1929.30
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Grating Support
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-114.2
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633.4
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3180.0
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B/H Mounting clam shells
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105.0
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827.5
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2366.5
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CCD Temperature monitor
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-116.75
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735.0
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1736.20
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Mid-box base plate
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0.0
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626.0
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1966.25
|
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Clamshell actuator
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67.30
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750.0
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2406.65
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These positions are shown in SR 8225 and the co-ordinates are
spacecraft.
The coordinate system is defined in RD 7.
6.9.6 Temperature Sensors
Type: PRT100 (integral with heater
pads)
6.9.7 Interface Temperatures
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Location
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Operating
|
Survival
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A0
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+10oC to +30oC
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-30oC to +40oC
|
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B0
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+10oC to +30oC
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-30oC to +40oC
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C0
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+10oC to +30oC
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-30oC to +40oC
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6.9.8 Properties of Outer Surfaces
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Emissivity
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Solar Absorptivity
|
IR Specularity
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Solar Specularity
|
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BOL
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EOL
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BOL
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EOL
|
|
|
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EIS MLI
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0.7
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0.7
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0.34
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0.34
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0.0
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0.0
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Radiators
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0.92
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0.92
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0.09
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0.16
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0.0
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0.0
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6.9.9 Operational Heaters
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Location
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Maximum power (W)
|
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Mirror Assembly
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2.5
|
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Grating Assembly
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2
|
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Mid Box
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1.5
|
|
|
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CCD (see note 1)
|
20
|
Note 1 – The CCD heater will be used to decontaminate the CCD and
will be used with other power systems switched down
6.9.10 Survival Heaters
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Location
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Power (W)
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Port #
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Mirror Assembly
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3.0
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1
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Grating Assembly
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2.0
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1
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Mid Box
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4.0
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2
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Readout Electronics
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1.0
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2
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CCD
|
5.0
|
3
|
7 INSTRUMENT CONTROL UNIT (ICU)
7.1 Mechanical
Interface
7.1.1 Mechanical details
Details of the mechanical interface of the ICU with the spacecraft are
shown in the drawing, ICU Interface A1 5275 300-3, see Appendix 3 (file
reference pal-5275-300-3).
7.1.2 Specification of attachment surfaces
The attachment point of the ICU is aluminium
Surface roughness = 1.6
μm CLA
(Centre-Line-Average)
7.1.3 Attachment Fastening
The four mounting holes have a diameter of 5.40 +0.30 - 0.05mm, to suit
10-32 UNF mounting bolts. The interface details are referenced in par
7.1.1.
The fastener torque is 35.7 - 39.1 kgf-cm (standard MS 16996 -
D/N)
7.2 Mass Properties
The mass of the ICU is given in section
5.1
7.2.1 Centre of Gravity
x = 163.697mm
y = 108.034mm
z = 56.006mm
These values are
calculated with respect to a reference datum identified on the drawing, ICU
Interface A1 5275 300-3, see Appendix 3 (file reference
pal-5275-300-3).
7.2.2 Moments of Inertia
Ixx = 0.1364 kg.m²
Iyy = 0.2551 kg.m²
Izz = 0.3448
kg.m²
These values are calculated with respect to a reference datum
identified on the drawing, ICU Interface A1 5275 300-3, see Appendix 3 (file
reference pal-5275-300-3).
7.3 Thermal
Interface
7.3.1 Attached Area
The base of the ICU will be machined flat but for the purpose of this
section it is assumed that each of the attachment bolts will provide TBD
cm
2 useful thermal contact area.
Total area of attachment of
the base of the ICU is 766.42
cm
2
7.3.2 Heat Dissipation Across Attachment
Points
|
Typical operating
|
Peak
|
Survival
|
|
Absolute W
|
21.3
|
25.3
|
0
|
|
Density W/cm2
|
0.028
|
0.033
|
0
|
7.3.3 Heat Capacity
The heat capacity of the ICU is TBD
W/K.
7.3.4 Acceptable Temperature Ranges
|
Operating
|
Survival
|
|
-20oC to +50oC
|
-30oC to +65oC
|
7.3.5 Positions of Temperature Measurement
The following sensors are provided by the spacecraft at the locations
indicated.
|
Sensor/Position
|
X
|
Y
|
Z
|
|
Near base datum
|
0
|
0
|
0
|
|
Top of front panel
|
295mm
|
121mm
|
0
|
These values are calculated with respect to a reference datum
identified on the drawing, ICU Interface A1 5275 300-3, see Appendix 3 (file
reference pal-5275-300-3).
7.3.6 Temperature Sensors
7.3.7 Interface Temperatures
|
Operating
|
Survival
|
|
-20oC to +50oC
|
-30oC to +65oC
|
7.3.8 Properties of Outer Surfaces
|
Emissivity
|
Solar Absorptivity
|
IR Specularity
|
Solar Specularity
|
|
BOL
|
EOL
|
BOL
|
EOL
|
|
|
|
EIS MLI
|
0.7
|
0.7
|
0.34
|
0.34
|
0.0
|
0.0
|
8 ELECTRICAL INTERFACES
There are five electrical interface between EIS-ICU (Interface Control
Unit) and Solar-B system components:
- ICU – DIST (Distributor)
- ICU – TCI-B (Telemetry Command Interface B)
- ICU – HKU (House Keeping Unit)
- ICU – HCE (Heater Control Electronics)
- ICU – MDP (Mission Data Processor)
There are two
electrical interfaces between EIS electrical components:
- ICU – ROE (Read Out Electronics)
- ICU – MHC (Mechanism and Heater
Controller)
EIS-STR
HCE
DIST
TCI-B
EIS-ICU
MDP
survival heater
#1-
#3
PIM
DHU
HKU
T-sensor
#1 -
#7
T-sensor
#8 -
#9
DC/DC
signal
line
primary power
on/off
CMD+TLM
28V primary
power line
secondary power
line
bus structure
- PIM
- PIM
- MHC
- ROE
Figure 3.1 Overview of electrical interfaces around
EIS-ICU
8.1 MDP—EIS-ICU Electrical Interface
The grounding scheme is shown in MSSL/SLB-EIS/DD002, see Appendix
5.
The scheme for cables and connectors is shown in MSSL/SLB-EIS/DD008, see
Appendix 7.
8.2 Overview of MDP— EIS-ICU Electrical Interface
EIS-ICU
MDP
Mission
Data
Serial
Command
Serial Status
Passive Bi-level
Status
Data
Enable
Clock
Data
Enable
Clock
Data
Enable
Clock
Data
Busy
4
ch
1
Figure 4.1 Overview of MDP-ICU electrical interface
1
line/component
MDM-25P
(MDP-11)
MDM-37P
(MDP-13)
There
are four types of electrical interfaces between MDP and EIS-ICU.
[1]
Bi-level status interface
Telemetry with bi-level status interface can be
transferred to MDP even when MDP-CPU is not running. There are four passive
bi-level status lines between MDP and ICU.
[2] Serial command
interface
Commands from the ground, DHU, and MDP to ICU are transferred to
ICU with this interface.
EIS HK status information is sent to
MDP with this interface.
[4] Mission data interface
Mission data
interface is a high-speed serial interface, which delivers science data from ICU
to MDP.
8.3 Passive Bi-level Interface
VDD: +5V
± TBD
VU1: CMOS 4049 / 4050Circuit
Type:
TB-E-RCircuit
Type:
TB-E-TC1: 220pF
± TBD
%R1: 27
kΩ
± TBD %R2: 10 k(
( TBD %EIS-ICUMDP VDDU1R1C1R2DATA 0DATA 4···Figure 4.2 Interface
circuit for passive bi-level status
interfaceU1R1C1VDDR2
- Point of definition for timing
chartSerial Command
Interface
Driver device:
HS-26C31Receiver device: HS-26C32
ENABLE
CLOCK
DATA
100Ω
±
10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
+
-
100Ω
±
10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
+
-
100Ω
±10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
+
-
EIS-ICU
MDP
insertion
of
choke coil
Figure 4.3-1 Interface circuit for serial command
interface
Clock signal comes only when serial command data
come out from MDP.
·
·
·
·
·
·
T1
T4
ENABLE
CLOCK
DATA
100%
50%
T3
T2
T: 1/f (f=62,500 Hz)
T1 :
T/2 ± 10
%
T2 : T
± 10
%
T3 : T/2
± 10
%
T4 : T/2
± 10
%
Figure 4.3-2 Timing chart for serial command
interface
Serial
Status Interface
point of definition for timing
chart
+
-
100Ω
±
10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
Receiver device:
HS-26C32
Driver device: HS-26C31
ENABLE
+
-
100Ω
±
10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
CLOCK
+
-
100Ω
±
10%
100Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
No
resistance
DATA
EIS-ICU
MDP
insertion
of
choke coil
?
Figure 4.4-1 Interface circuit for serial status
interface
Clock signal comes only when serial status data
come out from ICU.
·
·
·
·
·
·
T1
T4
ENABLE
CLOCK
DATA
100%
50%
T3
T2
T: 1/f (f=62,500 Hz)
T1 :
T/2 ± 10
%
T2 : T
± 10
%
T3 : T/2
± 10
%
T4 : T/2
± 10
%
Figure 4.4-2 Timing chart for serial status
interface
Mission
Data Interface
insertion
of
choke coil ?
position of definition
for timing
chart
ENABLE
CLOCK
DATA
+
-
33Ω
±
10%
33Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
100Ω
±
10%
+
-
33Ω
±
10%
33Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
100Ω
±
10%
+
-
33Ω
±
10%
33Ω
±
10%
4.7kΩ
±
10%
receiver
4.7kΩ
±
10%
driver
100Ω
±
10%
33Ω
±
10%
33Ω
±
10%
4.7kΩ
±
10%
driver
4.7kΩ
±
10%
-
+
100Ω
±
10%
BUSY
receiver
EIS-ICU
MDP
The grounding scheme is shown in MSSL/SLB-EIS/DD002, see
Appendix 5.
The scheme for cables and connectors is shown in
MSSL/SLB-EIS/DD008, see Appendix 7.
8.7 Electrical Pin-outs
8.8 System Voltage List
EIS takes 28V (max 30V) from the spacecraft and generates +5V, +8V, +15V,
+36V, -8V & -15V for sub-system use. +150V is generated in the MHC for drive
of the mirror fine positioner. No detector high voltage is necessary.
Transient voltages due to switching of inductive loads will be reduced
to less than 1V above the appropriate sub-system rail voltage.
Voltages
used by sub-systems:
ICU:
+5V, +15V, -8V, -15V.
MHC:
+8V, +15V,
-8V, -15V, -20V, +150V.
CAM:
+8V, +15V, +36V, -8V,
-15V.
8.9 Frequency List
Main power converter (ICU)
Operation at 200kHz (TBC). Total power is
about 30W.
Converter for fine mirror positioner driver (MHC)
Frequency
TBD. Total power about 2.4W.
Converter for 36V CAM bias
supply
Frequency TBD, phase locked to CAM clock. Total power about
800mW.
CAM Clock
In the range 16 - 32MHz.
ICU Processor
Clock
20MHz.
MHC Processor
Clock
9.816MHz.
9 POWER DISTRIBUTION
The power budget is given in paragraph 5.2.
Power distribution is
shown in MSSL/SLB-EIS/DD003.01, see Appendix 6.
10 SOFTWARE
INTERFACES
The ICU exchanges three types of data with the MDP. These are as
follows:
10.1 TC packets
Telecommand packets which consist of a command identifier (one byte)
followed by up to 132 bytes, as illustrated below:
|
CMD ID
|
Command Parameters
|
|
Command ID
(BC0)
|
Data Area
|
|
8 bits
|
Max. 132 bytes
|
The proposed EIS commanding structure is as follows:
|
CMD-IDs
|
Function
|
|
04 - 09
|
Memory dump (defined by MDP side - TBC)
|
|
E0 - EF
|
Memory uplink (defined by MDP side - TBC)
|
|
20 - 2F
|
Mode control commands
|
|
30 - 3F
|
PSU commands
|
|
40 - 4F
|
MHC commands
|
|
50 - 5F
|
CAM commands
|
|
60 - 6F
|
Flare operation commands
|
|
70 - 7F
|
Sequence table operations
|
|
80 – FF
|
Spares (excluding E0 – EF)
|
10.2 Status data
Status data packets (instrument HK) consist of a 4 bytes header followed
by up to 2 kbytes of status data. The general format is as
follows:
|
Header Area
|
Data Area
|
|
Data Type
|
Packet Size
|
Status Data
|
|
8 bits
|
24 bits
|
Max. ~2 kbytes
|
The following are the EIS status allocations:
Status type 1 -
ICU status
|
Header Area
|
Data Area
|
|
Data Type
|
Packet Size
|
EIS Status-1 Data
|
|
8 bits
|
24 bits
|
100 bytes
|
Status type 2 - ICU + PSU + CAM status
|
Header Area
|
Data Area
|
|
Data Type
|
Packet Size
|
EIS Status-1 Data
|
EIS Status-3 Data
|
|
8 bits
|
24 bits
|
100 bytes
|
150 bytes
|
Status type 3 - ICU + MHC status
|
Header Area
|
Data Area
|
|
Data Type
|
Packet Size
|
EIS Status-1 Data
|
EIS Status-3 Data
|
|
8 bits
|
24 bits
|
100 bytes
|
150 bytes
|
10.3 Mission data
Science data packets consist of a data header, followed by image data.
The maximum size of mission data packet is 256 kpixels (16 bit pixel). For
practical reasons a mission data packet is sent as a series of 4 kbyte
sub-packets, as illustrated below:
|
header
|
Image data 1
|
|
image data 2
|
|
image data 3
|
. . .
|
image data N
|
|
sub-packet
(e.g. 4Kbytes)
|
|
sub-packet
(e.g. 4Kbytes)
|
|
sub-packet
(e.g. 4Kbytes)
|
|
sub-packet
(e.g. < 4Kbytes)
|
Mission data parameters are still under discussion with
ISAS.
11 INSTRUMENT
MODES
The instrument modes are shown in the following diagram:
The following
table defines the instrument mode transition commands:
|
Command ID
|
Command Parameter
|
Mode
|
|
20
|
01
|
Standby
|
|
20
|
02
|
Manual
|
|
20
|
03
|
Auto
|
|
20
|
04
|
Emergency safe
|
|
20
|
05
|
Bake-out
|
|
20
|
06
|
Engineering
|
12 CONTAMINATION CONTROL
Details for the contamination control of the EIS instrument in its
assembly, integration and testing and commissioning phases are identified in the
Cleanliness Control Plan, RD 4.
12.1 Contamination
Tests
13 ENVIRONMENTAL
TESTS
13.1 Mechanical Tests
13.1.1 Test Matrix(sub-system)
|
Test/Model
|
MTM/TTM
|
PM
|
FM
|
|
Quasi-static load test 1
|
QT
|
N/A
|
N/A
|
|
Acoustic test 1
|
QT
|
N/A
|
PFT 2
|
|
Random vibration test 1
|
QT
|
N/A
|
PFT 2
|
|
Low frequency shock test
|
QT
|
N/A
|
N/A
|
|
Pyrotechnic shock test 1
|
TBD
|
N/A
|
N/A
|
1 Baseline is random vibration. Quasi static load test,
acoustic test and pyrotechnic shock test will only be performed if these loads
are found to be dominant in the design. Acoustic may be substituted for random,
if this is found to be dominant.
2 A protoflight test of either
acoustic or random will be performed, whichever is considered
dominant.
13.1.2 Test Matrix (Equipment within EIS)
13.1.3 Test Matrix(system)
|
Test/Model
|
MTM/TTM
|
PM
|
FM
|
|
Quasi-static load test
|
TBD
|
N/A
|
TBD
|
|
Acoustic test
|
QT
|
N/A
|
PFT
|
|
Random vibration test
|
TBD
|
N/A
|
TBD
|
|
Low frequency shock test
|
QT
|
N/A
|
PFT
|
|
Pyrotechnic shock test
|
QT
|
N/A
|
PFT
|
13.1.4 Test Levels
The test levels are defined in RD 6.
13.2 Thermal Vacuum Tests
13.2.1 Test Matrix
(System)
|
Test/Model
|
MTM/TTM
|
PM
|
FM
|
|
Thermal vacuum cycle
|
QT
|
N/A
|
AT
|
|
Thermal balance
|
QT
|
N/A
|
N/A
|
13.2.2 System Test
Levels
The temperature range for the QT is -30°C to +60°C
The
temperature range for the AT is -10°C to +50°C
The vacuum shall be
better than 10
-5 mm Hg
The Thermal Balance Test shall be
conducted at lower extremes of temperature than the QT (TBD) and for a duration
TBD.
Five cycles shall be performed for the thermal vacuum cycle test
(TBC).
13.2.3 Test Matrix (Sub-system)
|
Test/Model
|
MTM/TTM
|
PM
|
FM
|
|
Thermal vacuum cycle
|
N/A
|
N/A
|
AT
|
|
Thermal balance
|
N/A
|
N/A
|
TBD
|
13.2.4 Sub-system Test Levels
The temperature range for the AT is -10°C to +50°C
The vacuum
shall be better than 10
-5 mm Hg
Five cycles shall be performed
for the thermal vacuum cycle test
(TBC).
13.2.5 Test Matrix (Equipment within EIS)
13.3 EMC Tests
13.3.1 Test Matrix
|
Test/Model
|
MTM/TTM
|
PM
|
FM
|
|
Conducted and radiated emission and susceptibility
|
N/A
|
NONE
|
ALL
|
13.3.2 Test Levels
The test levels are identified in the Solar B Electrical Design Standard,
RD 8.
14 FAIRING ACCESS REQUIREMENTS FOR
CLAMSHELL
14.1 General
The EIS clamshell (CLM) is a vacuum compartment that protects the EIS
entrance filter from harmful environmental conditions (acoustics, contamination,
debris, humidity, air gusts, etc.) before and during Solar-B launch. The filter
is a very large thin aluminum filter and as such is quite fragile, an
inadvertent touch or rush of air will almost certainly destroy it. The only
safe way to launch such items is under a vacuum of <~1 torr. The CLM is a
special chamber for this purpose, and has two doors that open when the
instrument reaches orbit. It will remain under vacuum from the time of EIS
integration until safely deployed in
orbit.
14.2 Vacuum Requirement
The CLM is constructed to high vacuum standards and should be evacuable to
less than 0.001 Torr. Use of vacuum greases and lubricants are avoided and
organic materials are minimized. It is expected that the chamber can maintain a
vacuum of < 1 torr for up to 21 days. In all such vacuum systems without
active pumping, a gradual rise in pressure is expected with time due to
outgassing of material from internal surfaces, diffusion of gas through seals,
virtual leaks and real (but tiny) leaks. This means that the CLM will need to
be reattached to a pumping station periodically. Repumping should be done
whenever the pressure in the CLM approaches 1 torr. The frequency of this
operation must be determined during the commissioning of the CLM by plotting the
pressure versus time. Careful cleaning vacuum baking and leak testing will be
done to obtain the longest possible interval between pumping operations.
A
portable GSE pumping station has been designed to safely perform this operation
with the flight filters installed in the CLM. It can perform both evacuation
and backfilling operations and will follow the CLM throughout the EIS
development program and launch cycle. It is necessarily very slow in pumping
and backfilling since the filters would not survive any rush of air into or out
of the CLM. The pump connects to the CLM seal-off valve by a long copper tube
(flexible) and Swagelok connectors.
The seal-off valve is required to be an
integral part of the CLM to avoid having long pieces of tubing become part of
the CLM vacuum compartment. The CLM also provides a rugged support for the
valve body.
One or more vacuum gauges will be integral to the CLM so the
pressure can be monitored during the launch cycle. It is possible that one
gauge might be read by the MHC so the pressure could be found in the telemetry
whenever EIS is interrogated. A main vacuum gauge will be required that reads
out into an EGSE monitor. A [TBD] connector will be provided for this purpose.
Also integral to the CLM are a Photodiode and LED pair that allow checking the
filter for light tightness. Their EGSE unit can probably be combined with the
vacuum gauge EGSE.
14.3 Fairing Access
Using Astro-E as a model, fairing close-out is expected to occur on the
order of 12 days prior to launch. While this is less than the expected 21-day
hold time of the CLM, we must be prepared for situations where the hold time is
exceeded. Removing the fairing is much too difficult to contemplate, but a
hatch could provide the needed access to re-pump the CLM through the fairing.
The hatch opening would need to be on the order of 200 mm in diameter to permit
manual connection of the vacuum hose and operation of the seal-off valve.
The access hatch cover should be left open until 3 days before launch to
allow continuous pressure monitoring via the EGSE cable. A data logging
pressure monitor will be provided. A temporary hatch cover that passes the
cable can be used to protect the S/C. For reliable pressure measurements, the
same gauge, cable, and monitor should be used at all
times.
14.4 Access Hatch Location
The CLM is presently located at Z=+2334mm relative to the origin of the S/C
coordinate system. It has been suggested that the CLM seal-off valve be located
on the +Y facing surface of EIS at X=~105. In this case, the center of the
hatch should be at X=105, Y=~1100, Z= 2334 (S/C coordinates) in the fairing
skin. At this point there should be about 250mm clearance between the EIS +Y
panel and the fairing, an easy reach for an operator’s hand. If
necessary, small “finger wrenches” can be used in tight places.
These can be tied to the operator’s hand to prevent their being lost in
the fairing.
14.5 Expected Launch Pad Operations
There should be near continuous monitoring of the CLM pressure from the
time of the last pumpdown until three days (or less) prior to launch. The CLM
pressure will be plotted as a function of time and extrapolated to the end of
the launch window. Should the pressure be predicted to rise above a
predetermined limit, it will be necessary to re-pump the CLM.
This decision
should be made [TBD] days before scheduled launch to avoid last-minute chaos. A
final pumpdown within a week of the expected launch window should suffice in any
case. The launch preparation plan should include a critical decision point
where a pump/no pump decision is made that allows the pumping work to be carried
out with the minimum disruption of the rocket preparations.
The pumping
process will require bringing the GSE pumping station to the rocket building,
removing the hatch cover, connecting a vacuum hose, evacuating the hose,
verifying vacuum hose integrity, opening the CLM valve and pumping until the
desired pressure [TBD] is reached. The process is reversed at the end. The
pumping time depends on the size and length of the hose and can be baselined
during CLM development. Estimated working time should be about two hours to
start and two hours to close out after reaching the required pressure. The CLM
pressure should be monitored as long as possible after valve closure to verify
that a proper seal has been obtained. Two experienced vacuum operators should
be on hand to conduct this operation
safely.
15 NITROGEN PURGE
During integration and testing of the flight model with the spacecraft, the
EIS instrument will be required to be purged with clean, dry nitrogen gas
(specification TBC). This will be a common requirement with the other
instruments. Consideration should be given to a common supply with a combined
purge manifold for the instruments.
16 ACRONYMS
|
Acronym
|
Meaning
|
|
|
|
CCD
|
Charge Coupled Device
|
|
CLA
|
Centre-Line-Average
|
|
CLM
|
Clamshell
|
|
CNP
|
Connector Panel
|
|
EIS
|
EUV Imaging Spectrometer
|
|
FPA
|
Focal Plane Assembly
|
|
GA
|
General Assembly
|
|
ICU
|
Instrument Control Unit
|
|
MHC
|
Mechanism and Heater Controller
|
|
ROE
|
Read Out Electronics
|
|
TBC
|
To be confirmed
|
|
TBD
|
To be decided
|
