NAVAL RESEARCH LABORATORY SPACE SCIENCE DIVISION SOLAR PHYSICS BRANCH
Concept Study Report (CSR) Science/Technical/Management Volume for the Solar Orbiter Heliospheric Imager (SoloHI)
SSD-RPT-SOLOHI-0003 31 December 2011
Version 2.0
Approved By: Date: D. McMullin, SoloHI Project Engineer
Approved By: Date: S. Plunkett, SoloHI Project Manager
Approved By: Date: R. Howard, SoloHI Principal Investigator
DISTRIBUTION STATEMENT C: Distribution authorized to U.S. Government agencies and their contractors; Administrative or Operational Use. Other requests for this document shall be referred to Dr. R. Howard, Code 7660, Naval Research Laboratory, 4555 Overlook Avenue, S.W., Washington, D.C. 20375-5000.
4555 Overlook Avenue, S.W. Washington, D.C. 20375-5000 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003
RECORD OF CHANGES Revision Number Date Title or Brief Description Entered By
0.1 August 16, 2010 Draft CSR Submittal R. Howard
1.0 December 31, 2010 Final CSR Submittal R. Howard
2.0 December 31, 2011 Updated CSR Submittal R. Howard
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TABLE OF CONTENTS 1. EXECUTIVE SUMMARY ...... 1-1 1.1 Science Objectives ...... 1-1 1.2 Mission Overview ...... 1-1 1.3 Science Payload ...... 1-2 1.4 Key Spacecraft Characteristics ...... 1-2 1.5 Anticipated Launch Vehicle ...... 1-3 1.6 Mission Management ...... 1-3 2. SCIENCE INVESTIGATION DESCRIPTION ...... 2-1 2.1 Summary of Changes to Scientific Goals and Objectives ...... 2-1 2.2 Overview ...... 2-1 2.2.1 Scientific Goals and Objectives ...... 2-2 2.2.2 Value to NASA Goals & Objectives ...... 2-2 2.3 Scientific Goals and Objectives ...... 2-3 2.3.1 What are the source regions of the solar wind and heliospheric magnetic field? ...... 2-3 2.3.2 How do solar explosions produce energetic particle radiation that fills the heliosphere? ...... 2-5 2.3.3 How do solar transients drive heliospheric variability? ...... 2-7 2.3.4 What is the three-dimensional structure of the heliosphere? ...... 2-11 2.3.5 Analysis Techniques ...... 2-14 2.4 Observational Requirements ...... 2-14 2.4.1 SoloHI Signal to Noise ...... 2-14 2.4.2 Thomson Surface Considerations ...... 2-15 2.5 Science Data & Other Scientific Products ...... 2-16 2.5.1 Science Requirements Traceability Matrix (SRTM) ...... 2-16 2.5.2 Threshold Science Mission ...... 2-17 2.5.3 Descope Options ...... 2-18 2.5.4 Baseline Observing Programs ...... 2-27 2.6 Mission Operations ...... 2-28 2.6.1 Overview ...... 2-28 2.6.2 Science Operations ...... 2-28 2.6.3 Planning, Command Generation, and Telemetry Handling ...... 2-29 2.6.4 Mission Phases ...... 2-38 2.7 Data Reduction and Analysis ...... 2-39 2.7.1 Overview ...... 2-39 2.7.2 Data Flow ...... 2-39 2.7.3 Scientific Data Reduction ...... 2-40 2.7.4 Distribution ...... 2-40 2.7.5 Calibration...... 2-41 2.7.6 Data Analysis ...... 2-41
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2.7.7 Database Tools ...... 2-41 2.7.8 Computing Facilities ...... 2-41 2.8 Data Archive and Data Products ...... 2-41 2.8.1 Data Products ...... 2-42 2.9 Science Team ...... 2-42 3. TECHNICAL APPROACH ...... 3-1 3.1 Design Approach ...... 3-1 3.1.1 SoloHI Instrument ...... 3-1 3.1.1.1 SoloHI Design Modifications since the December 2010 CSR ...... 3-1 3.1.2 SoloHI Instrument Performance ...... 3-4 3.1.2.1 Telescope Scene Coverage ...... 3-2 3.1.2.1.1 Effective Area ...... 3-2 3.1.2.1.2 Spatial Resolution ...... 3-3 3.1.2.1.3 Straylight Rejection ...... 3-7 3.1.2.1.4 Modeling Coronal Focal Intensities for Orbital Positions of 0.28 and 0.7 AU ...... 3-8 3.1.2.1.5 SoloHI Photometric Analysis ...... 3-12 3.1.2.1.6 Strawman SoloHI Observing Program Characteristics ...... 3-13 3.1.3 SoloHI Design Heritage ...... 3-17 3.1.4 Baffle Design, Performance and Analysis ...... 3-18 3.1.4.1 Modifications to the Proposal Design ...... 3-22 3.1.5 Structure Design and Analysis ...... 3-22 3.1.5.1 Structure Characteristics ...... 3-22 3.1.5.2 Structure and Instrument Assembly ...... 3-23 3.1.5.3 Modifications to the Proposal Design ...... 3-24 3.1.5.4 Structural Analysis ...... 3-24 3.1.6 Preliminary Thermal Analysis ...... 3-25 3.1.6.1 Instrument Thermal Model Description ...... 3-26 3.1.6.2 Environment Thermal Model Description ...... 3-27 3.1.6.3 SIM Thermal Analysis Summary ...... 3-28 3.1.7 Electronics...... 3-29 3.1.7.1 Electronics Description ...... 3-29 3.1.7.2 Modifications to the Proposal Design ...... 3-34 3.1.7.3 Heritage ...... 3-35 3.1.8 Mechanisms ...... 3-35 3.1.8.1 Mechanisms Description ...... 3-35 3.1.8.2 Modifications to the Proposal Design ...... 3-35 3.1.8.3 Heritage ...... 3-35 3.1.9 Optical Design ...... 3-35 3.1.9.1 Optical Description ...... 3-35 3.1.9.2 Modifications since the 2010 CSR ...... 3-38
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3.1.9.3 Heritage ...... 3-38 3.1.10 Detector System ...... 3-38 3.1.10.1 Detector Description ...... 3-38 3.1.10.2 Testing of the SoloHI Minimal Prototype ...... 3-43 3.1.10.3 Modifications to the December 2010 CSR Design ...... 3-50 3.1.10.4 Heritage ...... 3-50 3.1.11 Focal Plane Assembly (FPA) ...... 3-50 3.1.11.1 FPA Description...... 3-50 3.1.11.2 Modifications to the Proposal Design ...... 3-51 3.1.11.3 Heritage ...... 3-51 3.2 Software Definition and Management ...... 3-51 3.2.1 Flight Software...... 3-51 3.2.1.1 Flight Software Architecture...... 3-51 3.2.1.2 Flight Software Requirements ...... 3-54 3.2.2 Ground Operations Software ...... 3-54 3.2.3 Software Test and Management...... 3-54 3.3 Technology Development Plan ...... 3-56 3.4 Spacecraft Interface and Accommodations ...... 3-56 3.4.1 Interface Definition ...... 3-56 3.4.1.1 Mechanical and Thermal Interfaces ...... 3-56 3.4.1.2 Electrical Interface ...... 3-56 3.4.1.3 Command and Data Handling Interface ...... 3-56 3.4.2 Interface Requirements and Accommodations ...... 3-56 3.4.2.1 Mass ...... 3-56 3.4.2.1.1 SoloHI Allocation ...... 3-56 3.4.2.1.2 Current Estimate ...... 3-57 3.4.2.1.3 Changes from the December 2010 CSR ...... 3-60 3.4.2.2 Operational Average Power ...... 3-60 3.4.2.2.1 SoloHI Allocation ...... 3-60 3.4.2.2.2 Current Estimate ...... 3-61 3.4.2.2.3 Changes from the December 2010 CSR ...... 3-63 3.4.2.3 Envelope ...... 3-64 3.4.2.3.1 SoloHI Allocation ...... 3-64 3.4.2.3.2 Current Estimate ...... 3-64 3.4.2.3.3 Changes from the December 2010 CSR ...... 3-64 3.4.3 Pointing Requirements ...... 3-65 3.4.4 Instrument Alignment Requirements ...... 3-65 3.4.5 Instrument Field of View ...... 3-66 3.4.6 Instrument Unobstructed Field of View ...... 3-66 3.4.7 On-Orbit Calibrations ...... 3-67 3.4.8 Cleanliness Requirements ...... 3-68
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3.5 Manufacturing, Integration, and Test ...... 3-68 3.5.1 Manufacturing, Integration, and Test for SoloHI Units and Assemblies ...... 3-68 3.5.1.1 SIM Structure ...... 3-69 3.5.1.2 Lens Barrel Assembly ...... 3-70 3.5.1.3 Straylight Baffle Assemblies ...... 3-70 3.5.1.4 APS Detector ...... 3-71 3.5.1.5 SoloHI Camera Electronics...... 3-71 3.5.1.6 Focal Plane Assembly (FPA) ...... 3-71 3.5.1.7 Baffle Cover Door Assembly ...... 3-72 3.5.1.8 SoloHI Power Supply ...... 3-72 3.5.2 Integration, Test, and Verification ...... 3-72 3.5.2.1 Science Payload Integration ...... 3-73 3.5.2.2 Verification and Validation Approach ...... 3-74 3.5.2.2.1 Requirement Level ...... 3-74 3.5.2.2.2 Hardware Levels of Assembly ...... 3-76 3.5.2.2.3 Hardware Models ...... 3-78 3.5.2.3 Verification Test Description ...... 3-79 3.5.2.3.1 Environmental Tests ...... 3-79 3.5.2.3.2 Functional Tests ...... 3-79 3.5.2.3.3 Performance Tests ...... 3-80 3.5.2.3.4 Characterization Tests ...... 3-80 3.5.2.3.5 Variance Tests ...... 3-80 3.5.2.4 Environmental Test Program ...... 3-81 3.5.2.4.1 SoloHI Flight Model ...... 3-81 3.5.2.4.2 SoloHI Structural Thermal Model ...... 3-81 3.5.2.4.3 Electronics...... 3-83 3.5.2.5 Functional, Performance, and Characterization Test Program ...... 3-84 3.5.2.6 Contamination Control and Cleanliness ...... 3-86 3.5.2.7 Facilities and Ground Support Equipment ...... 3-87 4. MANAGEMENT ...... 4-1 4.1 Management Processes ...... 4-1 4.1.1 Management Organization ...... 4-4 4.1.2 Decision-Making Process ...... 4-5 4.1.3 Management of Reserves and Margins ...... 4-5 4.1.4 Communications Process ...... 4-7 4.1.5 Staffing Plan...... 4-7 4.1.5.1 Principal Investigator ...... 4-7 4.1.5.2 Project Manager ...... 4-7 4.1.6 Systems Engineering Management Plan ...... 4-7 4.1.7 Management Approach ...... 4-8
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4.1.7.1 Project Management Control System ...... 4-8 4.1.7.2 Earned Value Management ...... 4-8 4.1.7.3 Requirements Process ...... 4-8 4.1.7.4 Mission Assurance Process ...... 4-9 4.1.7.5 Acquisition Process ...... 4-9 4.2 Schedule ...... 4-9 4.2.1 Critical Path ...... 4-10 4.3 Risk Management ...... 4-11 4.3.1 Top Risks and Mitigation...... 4-11 4.3.2 Risk Management Approach...... 4-11 4.3.3 Risk Management Tools ...... 4-11 4.3.4 Top Risks ...... 4-11 4.3.5 Managing New Technology Risks ...... 4-13 4.4 Government Furnished Property, Services, and Facilities ...... 4-14 4.5 Furnished Property, Services, and Facilities ...... 4-15 4.6 Reporting and Review Plan ...... 4-15 Appendix A. WORK BREAKDOWN STRUCTURE AND DICTIONARY ...... A-1 Appendix B. TRADE STUDIES SUMMARY ...... B-1 Appendix C. PHASE B WORK AGREEMENTS ...... C-1 Appendix D. PRODUCT ASSURANCE IMPLEMENTATION PLAN ...... D-1 Appendix E. SoloHI DETECTOR REQUIREMENTS, SELECTION, AND BACKUP ...... E-1 E.1 SoloHI Detector Driving Requirements ...... E-1 E.2 SoloHI Detector Trade Study Summary ...... E-2 E.3 SoloHI Detector Backup Plan ...... E-4 Appendix F. CONTAMINATION CONTROL PLAN...... F-1 Appendix G. LIST OF REFERENCES ...... G-1 Appendix H. LIST OF ACRONYMS...... H-1 Appendix I. INTEGRATED MASTER SCHEDULE ...... I-1 Appendix J. MASTER EQUIPMENT LIST (MEL) ...... J-1
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LIST OF FIGURES Figure 1-1. A false-color direct image from SECCHI/HI-1 (4°-24° from the Sun) showing the Milky Way, stars, the planet Mercury and solar coronal streamers. The vertical line is due to the shutterless operation on HI1, which will be avoided in SoloHI. The SoloHI FOV will cover 5.5°-45.5° from the Sun with higher spatial resolution ...... 1-1 Figure 1-2. Instrument Concept ...... 1-2 Figure 2-1. Velocity measurements of streamer blobs with LASCO. The SoloHI FOV during perihelion passages is also shown...... 2-4 Figure 2-2. Comparison of SECCHI/HI observation of solar wind structures (grey scale) with the model-derived location of the HCS (red surface). The meridional slice shows the model velocity but it is not discussed here...... 2-5 Figure 2-3. Left: A typical 3-part structured CME in LASCO/C2 on June 2, 1998. Right: A similar CME observed by SECCHI/HI1 on July 9, 2007 at a height of ~40 Rsun...... 2-7 Figure 2-4. Numerical Modeling of the Propagation of the 5/12/97 CME in the Heliosphere ...... 2-8 Figure 2-5. Schematic Representation of an Interplanetary CME ...... 2-9 Figure 2-6. Left panels: SECCHI/HI1 observations of the tail disconnection of Comet Encke. The images are shown in inverted brightness and are histogram-equalized to emphasize faint structures. Right panels: Running difference images. The vertical streaks are light from bright stars due to the shutterless operation of the SECCHI/HI cameras. The faint cloud approaching the comet is part of the CME front...... 2-10 Figure 2-7. (Top) Solar wind velocity and magnetic field superimposed on composite SOHO EIT & LASCO images contrasting the Sun and its wind at the minimum (first orbit) and maximum (second orbit) of solar activity. Solar Orbiter obtains 2 +/-30° scans per year...... 2-13 Figure 2-8. Expected contributions to the SoloHI signal as a function of elongation angle for perihelion (0.22 AU) and aphelion (0.88 AU). The 1- photon noise detection limit per pixel is shown for an exposure of 30 min at aphelion and 30 sec at perihelion. See Figure 3-5 for a revised figure with the correct perihelion distance of 0.28 AU...... 2-15 Figure 2-9. Sensitivity map of the Thomson scattering emission for the Solar Orbiter-Sun geometries at perihelion. Each number denotes a location in the orbit („2‟ is perihelion of 0.28 AU, while „1‟ and „3‟ are both at 0.34 AU). The 3 arches in each color mark the locus of the 5%, 50%, and 95% brightness integrals along the line of sight. The extent of the arches marks the SoloHI FOV. The dotted lines show the direction of the Parker Spiral for a 300 km/s wind...... 2-16 Figure 2-10. SoloHI Requirements Flowdown from Level 1 to Level 3 ...... 2-17 Figure 2-11. Spacecraft Distance from Sun During Solar Orbiter Mission ...... 2-31 Figure 2-12. Spacecraft Distance from Sun and Solar Latitude During Solar Orbiter Mission ...... 2-31 Figure 2-13. SoloHI Concept of Operations ...... 2-37 Figure 3-1. Physical block diagram. Diagram shows the SoloHI components. The SoloHI consists of the SoloHI Instrument Module (SIM) and the SoloHI Power Supply Box (SPS)...... 3-2
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Figure 3-2. SoloHI instrument module optomechanical layout. The SIM design is compact with straightforward optomechanical interfaces...... 3-3 Figure 3-3. Scene Image Projected on the SoloHI Detector ...... 3-2 Figure 3-4. F and K Equatorial Corona Brightness Estimate at Various Spacecraft-Sun Distances with Equivalent Photon Noise Shown for 1 and 10s Exposure Times ...... 3-11 Figure 3-5. Count rates (electrons/s/pixel) at the instrument focal plane for the corresponding F and K Equatorial Corona. Curves follow the same legend given in the previous figure...... 3-11 Figure 3-6. The SoloHI instrument Concept and Mechanical Layout is Similar to SECCHI/HI ...... 3-18 Figure 3-7. Baffle system side view schematic. Dimensions show general scale but are not for reference. The shadow line shows the worst case edge of the sun at 2.05 degrees. F1 to F4 shade the lens aperture from diffracted radiation scattered at the edge of the heat shield. Interior and peripheral baffles protect the instrument from scattered radiation from the surrounding spacecraft ...... 3-19 Figure 3-8. Baffle system top view schematic. Placement of interior baffles and front baffles shown. Dimensions show general scale but are not for reference. Details of side wall baffle and lens cavity baffle arrangement are not shown ...... 3-19 Figure 3-9. Calculated irradiance at the lens entrance aperture. The calculation was done for the perihelion case, worst case off-point with a nominal heat shield position. Dashed lines show the lens entrance aperture ...... 3-20 Figure 3-10. Calculated irradiance at the lens entrance aperture. The calculation was done for the perihelion case, worst case off-point with a worst case heat shield position. Dashed lines show the lens entrance aperture ...... 3-20 Figure 3-11. Solar array and RPW antenna geometry relatively to SoloHI and the heat shield. Figure generated from Zemax non-sequential ray trace analysis ...... 3-21 Figure 3-12. RPW antenna illuminating the SoloHI instrument (only the brightest rays shown). Figure generated from Zemax non-sequential ray trace analysis ...... 3-21 Figure 3-13. Solar array illumination of the SoloHI instrument baffle (brightest rays only). Figure generated from Zemax non-sequential ray trace analysis ...... 3-22 Figure 3-14. Instrument Exploded View ...... 3-23 Figure 3-15. Instrument Internal Baffle Box ...... 3-24 Figure 3-16. Structural Deformation Plot Showing Primary Lateral Z Frequency ...... 3-25 Figure 3-17. SoloHI Thermal Model ...... 3-26 Figure 3-18. SoloHI Thermo-Optical Properties ...... 3-26 Figure 3-19. System Electrical Block Diagram ...... 3-30 Figure 3-20. Camera Electronics Card (CEC) Architecture ...... 3-31 Figure 3-21. SoloHI Camera Electronics Breadboard (CEBB) ...... 3-31 Figure 3-22. Processor Card (PRC) Architecture ...... 3-32 Figure 3-23. SoloHI Power Supply Cards ...... 3-33 Figure 3-24. SPS Enclosure ...... 3-33 Figure 3-25. SoloHI Lens Layout ...... 3-36 Figure 3-26. SoloHI telescope field of view ...... 3-37
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Figure 3-27. Lens Spot Diagrams as a Function of Field Angle ...... 3-37 Figure 3-28. Off-Axis lens stray light performance. The three curves show the sensitivity to the lens barrel geometry. The orange curve is the SoloHI baseline design ...... 3-38 Figure 3-29. 6T PPD Pixel Design Schematic ...... 3-39 Figure 3-30. SoloHI APS block diagram. The individual pixel controls are addressable on a row by row basis ...... 3-40 Figure 3-31. Stitched reticule geometry. This figure shows how stitching is used to construct a large scale array ...... 3-41 Figure 3-32. SoloHI APS detector map showing the stitching blocks ...... 3-41 Figure 3-33. Stitching patterns. Preliminary results from the stitching study show that 5 stitching patterns (P1 to P5) are required to stitch the SoloHI array ...... 3-42 Figure 3-34. Preliminary 4kx4k device wafer layout with 1kx1k test imagers ...... 3-43 Figure 3-35. SoloHI minimal device performance as measured by Jim Janesick. This shows a full well of 35k and 154k electrons with noise of ~4.8 and ~25 electrons respectively ...... 3-44 Figure 3-36. NRL CEBB custom camera built by Silver Engineering (top). NRL test chamber with optical port being used for minimal device testing (bottom) ...... 3-45 Figure 3-37. Photon transfer curve for the Single Pixel readout of the Sandbox VI Detector ...... 3-46 Figure 3-38. Read Noise for a Row of Pixels for the Sandbox VI Detector at High Gain using the Camera Electronics Breadboard (CEBB) ...... 3-46 Figure 3-39. Linear Full Well for a Row of Pixels for the Sandbox VI Detector at High Gain using the Camera Electronics Breadboard (CEBB) ...... 3-47 Figure 3-40. Dark Current Rate for Various Sandbox VI Devices over the Temperature Range of -73 to +7 deg C and over the Voltage Range of 3.6 to 4.5 V ...... 3-47 Figure 3-41. Dark Images for Various Sandbox VI Devices over the Temperature Range of -73 to -33 deg C and over the Voltage Range of 3.6 to 4.5 V ...... 3-48 Figure 3-42. Relative Wavelength Response for Various Sandbox VI Devices over the Bandpass of 400 to 1100 nm and over the Voltage Range of 3.6 to 4.5 V, measured by GSFC Detector Branch. Quantum Efficiency is roughly 25% over the 490 to 740 nm bandpass. Final calibration has not been applied yet ...... 3-48 Figure 3-43. Linear Response to a High Signal for Sandbox VI Devices at High Gain, measured by GSFC Detector Branch ...... 3-49 Figure 3-44. Linear Response to a Low Signal for Sandbox VI Devices at High Gain, measured by GSFC Detector Branch ...... 3-49 Figure 3-45. FPA design concept. The figure shows the side mounting of the FPA in the SoloHI structure. Preliminary sizing of the Kovar mount and detector readout board are shown ...... 3-50 Figure 3-46. SoloHI focal plane assembly exploded view. The SoloHI focal plane assembly uses a minimal number of parts. The 3 boss detector mounting surface is not shown ...... 3-51 Figure 3-47. SoloHI Flight Software Architecture ...... 3-51 Figure 3-48. Flight Software Module Interaction with Electronics and Software Bus ...... 3-53
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Figure 3-49. SIM Front Baffle to Heat Shield Leading Edge Nominal Location and Alignment Tolerance ...... 3-66 Figure 3-50. SIM Unobstructed FOV Boundary Elevation Over the Azimuth Range ...... 3-68 Figure 3-51. Development Model Baffle Assembly ...... 3-70 Figure 3-52. SoloHI Assembly, Integration, and Test Flow ...... 3-73 Figure 3-53. SoloHI Requirements Flowdown from Level 1 to Level 3 ...... 3-75 Figure 3-54. SoloHI Requirements Flowdown from Level 3 to Level 5 ...... 3-76 Figure 3-55. Hardware Levels of Assembly Hierarchy ...... 3-77 Figure 3-56. SoloHI Hardware Assembly Levels (Expanded to Unit Level) ...... 3-78 Figure 4-1. Our Streamlined Organizational Structure Provides Clear and Simple Lines of Project Accountability and Communication ...... 4-2 Figure 4-2. A Systematic Decision-Making Process Supports the Use of Margins and Reserves ...... 4-6 Figure 4-3. SoloHI Risk Matrix ...... 4-13
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LIST OF TABLES Table 1-1. Instrument Resources and Interfaces ...... 1-3 Table 2-1. SoloHI Threshold Mission Measurement Requirements Comparison to STEREO HI ...... 2-18 Table 2-2. SoloHI Science Requirements Traceability Matrix (1 of 4) ...... 2-19 Table 2-3. SoloHI Level 1 Science Measurement Requirements for the Baseline and Threshold Science Mission ...... 2-23 Table 2-4. SoloHI Level 3 Science Measurement Requirements for the Baseline and Threshold Science Mission (1 of 2) ...... 2-24 Table 2-5. SoloHI Program Descope Options ...... 2-26 Table 2-6. Radial Scene Coverage for SoloHI Images for Each Orbital Region Boundary ...... 2-30 Table 2-7. Mission Phases and Individual Orbit Timeline During Baseline Solar Orbiter Mission with a Jan 1, 2017 Launch Date ...... 2-32 Table 2-8. SoloHI Baseline Observing Program for Orbit I of Solar Orbiter Mission ...... 2-33 Table 2-9. SoloHI Baseline Observing Program for Orbit II of Solar Orbiter Mission ...... 2-34 Table 2-10. SoloHI Data Volume Estimate for Orbit I Baseline Observing Program ...... 2-35 Table 2-11. SoloHI Data Volume Estimate for Orbit II Baseline Observing Program ...... 2-36 Table 2-12. Nominal Operational Strategy for the SoloHI at Each of the Mission Phases ...... 2-38 Table 2-13. SoloHI Data Products and Formats ...... 2-42 Table 2-14. A World-Class Science Team with Expertise in Instrument Development and Solar Science ...... 2-43 Table 3-1. SoloHI Nominal Design Characteristics Table ...... 3-3 Table 3-2. SoloHI Instrument Performance Requirements and Capabilities ...... 3-5 Table 3-3. SoloHI Instrument Performance Requirement Margin ...... 3-1 Table 3-4. SoloHI Telescope Effective Area Budget Requirements and Capabilities ...... 3-2 Table 3-5. SoloHI Spatial Resolution Budget Requirements and Capabilities for Full Frame and Inner FOV Subframe Images ...... 3-4 Table 3-6. Nominal Spatial Resolution Across the Image Field for SoloHI Full Frame, Inner FOV Subframe and Radial Swath Subframe Images ...... 3-5 Table 3-7. SoloHI Image Spatial Resolution Budget ...... 3-6 Table 3-8. SoloHI Straylight Rejection Requirement as a Fraction of the Coronal Brightness ...... 3-7 Table 3-9. Lens Barrel Straylight Rejection Requirement for Aperture Incident Angle Ranges ...... 3-7 Table 3-10. SoloHI Straylight Rejection Requirements and Capabilities for Full Frame Images ...... 3-9 Table 3-11. SoloHI Straylight Rejection Modeled Estimate at Detector Inner (Top Table) and Outer (Bottom Table) FOV ...... 3-10 Table 3-12. SoloHI Minimum Integration Times for Perihelion and Near Perihelion Observing Cases...... 3-14 Table 3-13. SoloHI Minimum Integration Times for Far Perihelion, North and South Observing Cases...... 3-15
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Table 3-14. SoloHI Baseline Observing Program Characteristics...... 3-16 Table 3-15. Stray Light Analysis Summary ...... 3-22 Table 3-16. Primary Modes for the SoloHI Instrument Module on Instrument Mounts ...... 3-25 Table 3-17. SoloHi Optical Properties ...... 3-27 Table 3-18. SoloHI Thermal Case Assumptions ...... 3-28 Table 3-19. SoloHI Power Dissipations ...... 3-28 Table 3-20. SoloHI Temperature Predictions ...... 3-28 Table 3-21. SoloHI Heater Power Estimates ...... 3-29 Table 3-22. Phase A RTC Trade Study ...... 3-34 Table 3-23. SoloHI Lens Prescription ...... 3-36 Table 3-24. Selected APS Characteristics ...... 3-39 Table 3-25. Summary of Current Sandbox VI Performance Test Results ...... 3-44 Table 3-26. SoloHI Flight Software Modules...... 3-52 Table 3-27. SoloHI Flight Software Tasks ...... 3-53 Table 3-28. LEON3 FT Image Compression Summary ...... 3-54 Table 3-29. Flight/Ground Software Build Strategy ...... 3-55 Table 3-30. Mass Breakdown for the SoloHI Science Payload ...... 3-59 Table 3-31. SoloHI Science Payload Mass Change since the 12/10 CSR ...... 3-60 Table 3-32. Average Power Breakdown for the SoloHI Science Payload ...... 3-63 Table 3-33. SoloHI Science Payload Operational Average Power Change since the December 2010 CSR...... 3-64 Table 3-34. SIM Unobstructed FOV Definition ...... 3-67 Table 3-35. Providers for SoloHI Units and Assemblies ...... 3-69 Table 3-36. Environmental Test Matrix for SoloHI Hardware Levels of Assembly at Unit Level or Higher ...... 3-82 Table 3-37. Cycles, Failure-Free Hours and Burn-in Hours for SoloHI Electronics ...... 3-84 Table 3-38. SoloHI Verification and Characterization Test Matrix ...... 3-85 Table 3-39. Variance Test Matrix for SoloHI Flight Hardware ...... 3-86 Table 4-1. The PI has Implemented Well-Defined Roles, Responsibilities, and Commitments With an Experienced Institution, the NRL, to Define, Implement, and Execute the SoloHI Project ...... 4-3 Table 4-2. SoloHI Key Team Members ...... 4-3 Table 4-3. SoloHI Key Milestones ...... 4-10 Table 4-4. SoloHI Risk Table Based on Risk Exposure ...... 4-12
Table A-1. Work Breakdown Structure and Dictionary ...... A-1 Table B-1. Trade Studies Summary ...... B-1 Table E-1. Detector Driving Requirements that Led to the APS Detector Selection ...... E-1 Table E-2. SoloHI Detector Candidate Characteristics ...... E-2 Table E-3. Detector Candidate Programmatic and Technical Accommodation Requirements ...... E-3
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Table E-4. Projected Budget Impact for SoloHI Detector Contingency Plans ...... E-6 Table J-1. Basis of Estimate Definitions ...... J-1 Table J-2. SoloHI Master Equipment List ...... J-2
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1. EXECUTIVE SUMMARY 1.1 Science Objectives The Solar Orbiter Heliospheric Imager (SoloHI) will perform remote observations of the Thomson-scattered light from the solar wind plasma from the inner corona to the orbit of Venus, with unprecedented resolution and cadence. SoloHI, built upon the successful SECCHI Heliospheric Imager (HI) aboard the STEREO mission, addresses the four key questions in the 2009 Solar Orbiter Assessment Study Report: What are the origins of the solar wind streams and the heliospheric magnetic field? What are the sources, acceleration mechanisms, and transport processes of solar energetic particles? How do coronal mass ejections evolve in the inner heliosphere? How does the solar dynamo work and drive connections between the Sun and the heliosphere?
SoloHI will be able to image the solar wind as it travels away from the Sun and impinges on the Solar Orbiter and other inner heliospheric probes, including Solar Probe Plus. This will enable us to determine the relation of the solar wind structures seen by the different spacecraft. Figure 1-1 demonstrates the anticipated SoloHI imaging performance using an actual observation from SECCHI/HI1. SoloHI will have twice the field of view (FOV) of the existing SECCHI/HI1 instrument with improved spatial resolution, and comparable signal-to-noise ratio. As the Solar Orbiter approaches the Sun, the resolution will increase relative to its resolution at 1 AU and the FOV will decrease relative to 1 AU. For example, at perihelion the SoloHI will have Figure 1-1. A false-color direct image from the same effective resolution as the SOHO SECCHI/HI-1 (4°-24° from the Sun) showing LASCO/C2 coronagraph, with the same field the Milky Way, stars, the planet Mercury and of view as LASCO/C3, but still with excellent solar coronal streamers. The vertical line is due signal-to-noise ratio. to the shutterless operation on HI1, which will be avoided in SoloHI. The SoloHI FOV will 1.2 Mission Overview cover 5.5°-45.5° from the Sun with higher The Solar Orbiter is a three-axis stabilized spatial resolution spacecraft, carrying both in-situ and remote sensing instruments. Using gravity assists from Earth and Venus, Solar Orbiter will be placed into an elliptic orbit, with a perihelion of about 0.28 AU and an aphelion of about 0.88 AU. The
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Venus gravity assists will also raise the inclination of the orbital plane to heliolatitudes of about 30º by the end of the 7-year nominal mission. 1.3 Science Payload The Solar Orbiter payload has been defined, consisting of both in-situ and remote sensing instruments. The in-situ instruments will measure the kinetic properties and composition of the solar wind plasma, the solar wind magnetic field and plasma waves and the solar energetic particles. The remote sensing instruments will image the solar magnetic field and velocity fields, the EUV and visible light corona, and include EUV and X-ray spectrographs, and a heliospheric imager.
This SoloHI Concept Study Report (CSR) addresses the heliospheric imager. It is a single, white light telescope of 20º half angle with the inner limit of the field of view at an elongation of 5.5º from Sun center. The key aspects of the SECCHI/HI instrument design was to reject light from the solar disk to an acceptable level and to accumulate sufficient exposure to have a good signal to noise ratio even Figure 1-2. Instrument Concept when observing the weak solar wind plasma. For the Solar Orbiter mission there are two added complexities – the stray light scattered by the solar array at the rear of the instrument and the effect of energetic particles, particularly protons and neutrons. To address these additional requirements, the baffle design has been optimized to reject the reflected light and an active pixel sensor (APS) detector is used to avoid the charge transfer efficiency problems of an irradiated CCD. Figure 1-2 shows the SoloHI instrument concept. It consists of a simple box structure with a protective lid, various baffles, a simple lens system, an APS imaging detector, camera electronics and a passive radiator. 1.4 Key Spacecraft Characteristics The Solar Orbiter Spacecraft is three-axis stabilized with one axis sun centered. The design includes a heat shield extending on all sides of the front face as a means to protect the spacecraft from the intense solar flux at perihelion. The remote sensing instruments requiring a solar view shall have an opening in the heat shield. The in-situ and SoloHI instruments shall mainly be in the shade of the sunshield. The pointing control system is based on reaction wheels and chemical propulsion thrusters. The solar arrays deploy after launch and have a single axis articulation. The RF subsystem has one HGA and one MGA for science telemetry downlink.
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ESA has completed a pre-phase A study focused on developing a design to reject the heat generated during the perihelion passage and to accommodate a representative set of instruments. Much of it utilizes designs from their Bepi-Colombo mission to Mercury, which is scheduled to launch before Solar Orbiter. The design will be matured during the upcoming Phase A/B. Table 1-1 provides the basic SoloHI instrument resources and spacecraft interfaces. 1.5 Anticipated Launch Vehicle NASA will provide the launch vehicle which is to be launched from the Kennedy Space Center, Cape Canaveral, Florida. 1.6 Mission Management Solar Orbiter is part of the joint ESA-NASA International Living With A Star (ILWS) program, which includes the Solar Orbiter and the Solar Probe Plus spacecraft. ESA is responsible for the Solar Orbiter spacecraft, integrating the instruments and mission operations. NASA Goddard Space Flight Center manages the U.S. hardware efforts, provides management interfaces with the ESA Solar Orbiter project office and provides the launch services. NRL is the SoloHI PI institution and will deliver the assembled, calibrated instrument to NASA. NRL is responsible for the structure, the optical system, APS detector, electronics, software, program management and thermal and structural analyses. NRL will conduct calibration and environmental testing, conduct the tests at observatory level and then will operate the instrument after launch. Table 1-1. Instrument Resources and Interfaces Spacecraft Science Instrument Element SoloHI Design with Reserves Subsystem SIM Mass 9.65 kg (CBE) SPS Mass 2.83 kg (CBE) SIM to SPS Harness Mass 0.25 kg (CBE) SoloHI Total Mass 12.74 kg (CBE); 15.50 kg (Allocation) Mechanical SIM/BracketsEnvelope (Door 63.0 cm Xopt x 21.0 cm Zopt x 42.0 cm Closed) Yopt 16.85 cm Xopt x 17.0 cm Zopt x 23.0 cm SPS Envelope Yopt Field of View -40 deg Radial x 40 deg Transverse SoloHI Operational Average Power 9.44 W (CBE); 12.0 W (Allocation) Operational Heaters 2.4 W (Out of Ecliptic) Power Survival Power 6.9 W (CBE) Primary Power +28 Vdc +1/-2 Vdc Data Interfaces Spacewire Downlink Data Rate 20.5 kbps Downlink Data Volume per Orbit 53.136 Gbits (w/Housekeeping, CCSDS) C&DH/RF Uplink Daily Data Volume 60 kbits, Avg Command Data w/CCSDS One Shot Door Operation +28 Vdc Wax Actuator
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2. SCIENCE INVESTIGATION DESCRIPTION 2.1 Summary of Changes to Scientific Goals and Objectives Since the submission of the SoloHI Concept Study Report (CSR) in December 2010, the Solar Orbiter Collaboration Project was instructed by NASA HQ to continue in the Phase A Extension until December 31, 2011. During this time, the SoloHI team was instructed to investigate cost reduction options and submit a white paper to the Solar Orbiter Collaboration Project. The Project approved two design changes; namely, the elimination of the flight calibration LEDs and the elimination of the Cameralink interface. In addition the SoloHI Control Electronics were moved under the instrument from the spacecraft interior. These changes are described in detail in Section 3. While these changes affect the instrument design, they do not drive any changes in the scientific goals and objectives of the SoloHI Science Investigation. Over the last year, we refined further the SoloHI observing programs as the orbits were better defined by ESA. Therefore, the SoloHI Science Investigation has been updated to reflect these small changes. The Science Requirements Traceability Matrix (SRTM) remains the same as in the 2010 CSR.
2.2 Overview The Earth and other bodies in the solar system exist within the tenuous outer atmosphere of our nearest star, the Sun, in a volume of space known as the heliosphere. The newly-integrated science of heliophysics seeks to understand how the structure and dynamics of the solar system are controlled by the outflow of plasma and magnetic field from the Sun. A key objective of this new science is to fully characterize, and ultimately be able to predict, the space environment in the heliosphere. The Solar Orbiter program will combine multipoint in-situ measurements with high-resolution remote-sensing observations of the Sun from Solar Orbiter in a systemic approach to resolve fundamental science problems needed to achieve this objective. These unsolved problems include the sources of the solar wind, the causes of eruptive releases of plasma and magnetic field from the Sun known as coronal mass ejections (CMEs), the evolution of CMEs and their interaction with the ambient solar wind flow, and the origins, acceleration mechanisms and transport of solar energetic particles that may be hazardous to both human explorers and robotic spacecraft that operate in the highly variable environment outside the protective cocoon of Earth‟s atmosphere. The SoloHI will provide the crucial link between the remotely-sensed and in-situ observations from Solar Orbiter and from the upcoming Solar Probe Plus (SPP) and Bepi-Colombo missions (referred to as Inner Heliospheric Probes (IHP), hereafter) that is needed to address these fundamental problems. SoloHI will image both the quasi-steady flow and transient disturbances in the solar wind over a wide field of view that will encompass much of the orbit of the Sentinels spacecraft, by observing visible sunlight scattered by electrons in the solar wind. The SoloHI FOV is centered on the ecliptic plane but is offset from the Sun, and covers a range of elongation angles from 5.5° to 45.5°. SoloHI will provide synoptic images of the inner heliosphere with good spatial resolution that will maximize the scientific return of both the Solar Orbiter and the IHP programs. The observations will image both co-rotating and transient CME structures as these structures propagate through the heliosphere and ultimately pass over one or more of the IHPs. Thus, SoloHI will provide a broad context for interpreting the in-situ measurements to be made by Solar Orbiter and the IHPs.
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2.2.1 Scientific Goals and Objectives A half century of observations and theoretical investigations has provided an increasingly detailed understanding of the Sun and the solar wind. However, progress on several basic questions has been hampered by the paucity of coordinated remotely-sensed and in-situ observations from multiple perspectives, including out of the ecliptic plane, close enough to the Sun that fundamental processes involved in the generation of the solar wind and the acceleration of solar energetic particles can be understood. The SoloHI science investigation has four science objectives that have been identified by the NASA and ESA Joint Science and Technology Definition Team (JSTDT) and the 2009 Solar Orbiter Assessment Study Report (ASR) as the focus of the Solar Orbiter program. These objectives will be addressed by answering the following science questions: How and where does the solar wind plasma and magnetic field originate in the corona? – What are the source regions of the solar wind and heliospheric magnetic field? – What mechanisms heat and accelerate the solar wind? – What are the sources of solar wind turbulence and how does it evolve? How do solar transients drive heliospheric variability? – How do CMEs evolve through the corona and inner heliosphere? – How do CMEs contribute to the solar magnetic flux and helicity balance? – How and where do shocks form in the corona? How do solar explosions produce energetic particle radiation that fills the heliosphere? – How and where are energetic particles accelerated at the Sun? – How are solar energetic particles released from their sources and distributed in space and time? – What are the seed populations for energetic particles? What is the three-dimensional structure of the heliosphere? – What is the three-dimensional structure and extent of streamers and CMEs? – How are variations in the solar wind linked to the Sun at all latitudes? – What are the sources and properties of dust in the inner heliosphere, and do Sun-grazing comets contribute to the dust? These objectives are discussed in detail in the following sections. For some of the objectives, SoloHI observations are required in order to answer the underlying questions, while for others, SoloHI observations will provide the context required to interpret other Solar Orbiter measurements. 2.2.2 Value to NASA Goals & Objectives The SoloHI science objectives are directly relevant to the NASA Strategic Goal 3.2 to „understand the Sun and its effects on the Earth and the solar system‟, and will specifically
2-2 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 address several Heliophysics research objectives and focus areas. The SoloHI observations will be critical for the joint science between remote and in-situ instruments on Solar Orbiter that forms the core of the Solar Orbiter program, and would also significantly enhance the science return from the upcoming SPP mission. SoloHI will likely be the only heliospheric imager capable of providing the necessary observations at the time of the Solar Orbiter, SPP, and Bepi- Colombo missions.
2.3 Scientific Goals and Objectives The Solar Orbiter mission design is quite complex, involving perihelion and aphelion passages which do not always occur at the same heliocentric distance and/or with the same inclination relative to the ecliptic. For an imaging instrument such as SoloHI, these orbital variations imply both a changing field of view, in terms of heliospheric coverage, and varying spatial resolution, thus determining the optimum observing strategy and science focus for a given objective (e.g., shock formation versus turbulence studies). To appreciate the flexibility and performance of the proposed SoloHI instrument and to facilitate comparison with similar instruments at Earth orbit, we express the FOV and spatial resolution in terms of their 1 AU equivalent quantities using the symbol, AUeq, for shorthand.
2.3.1 What are the source regions of the solar wind and heliospheric magnetic field?
Key Science Questions What are the source regions of the solar wind and heliospheric magnetic field? What mechanisms heat and accelerate the solar wind? What are the sources of solar wind turbulence and how does it evolve?
White-light imaging with the LASCO coronagraphs has revealed a variety of unexpected dynamical phenomena in the outer corona, including plasma blobs that are ejected continually from the cusps of streamers (Sheeley et al. 1997; Wang et al. 1999b, 2006), ray-like structures pervading the streamer belt (Thernisien & Howard, 2006), and swarms of small-scale inflows (Wang et al. 1999a; Sheeley & Wang 2001; Sheeley et al. 2001) that occur during times of high solar activity. Some of these phenomena can be interpreted as different manifestations of field- line interchange reconnection, in which plasma and magnetic flux are exchanged between closed and open field regions of the corona. Such reconnection processes have a bearing on questions as diverse as the formation and evolution of the heliospheric plasma/current sheet, the origin of the slow solar wind, the heliospheric magnetic flux budget, the solar-cycle evolution of the coronal field, and the rigid rotation of coronal holes. While the reconnection sites will be observed with the Solar Orbiter coronagraph and EUV imager in the inner corona, SoloHI observations are essential for measuring the outer corona and heliospheric signatures of these events. During perihelion, SoloHI will be able to trace the streamer blobs, formed in the Solar Orbiter coronagraph FOV, to much greater heights and with better resolution than is possible with the LASCO or even the SECCHI coronagraphs. Figure 2-1, Table 2-9, and Table 2-10 show that the proposed observing programs (A1.1, A1.2) are
2-3 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 similar to the LASCO synoptic program (24-45 min cadence), and will be sufficient to meet this objective. The large uninterrupted FOV of SoloHI enables more accurate velocity and mass measurements, compared to LASCO or SECCHI, and the increased resolution and sensitivity of SoloHI will reduce the scatter in the outer velocity measurement shown in Figure 2-1. SoloHI will measure the velocity and acceleration profile of the transient slow solar wind flows and also the mass of this component.
As the blobs propagate into the heliosphere, they expand and merge with each other forming large scale structures. The recent analysis of SECCHI observations has clearly demonstrated that both fast and slow speed streams can be imaged and their kinematics measured with the SECCHI Heliospheric Imagers (HI) from near the Sun to 1 AU (Sheeley et al. 2008). The HI1 images show that the density compression ahead of a co-rotating interaction region (CIR) starts to develop as early as 0.3 AU in some cases. These observations suggest that imaging observations of solar wind structures are essential to properly interpret in-situ measurements because interactions among the various solar wind structures can start close to the Sun. SoloHI will provide these crucial observations over a wide field of view, including inner heliospheric probes, during some of its passages (Table 2-9 and Table 2-10). Figure 2-1. Velocity measurements of streamer blobs with LASCO. The SoloHI will investigate the structures that comprise the SoloHI FOV during perihelion heliospheric plasma sheet (HPS) and their solar origins, passages is also shown. and the relation of the HPS to the heliospheric current sheet (HCS). For this, SoloHI will image the extension of streamer structures far into the heliosphere and compare their measured location and densities to in-situ measurements and models. Again, the SECCHI/HI observations have shown that this is possible. In Figure 2-2, taken from Vourlidas & Riley (2007), the location of the HCS, based on an MHD simulation, is projected onto a 2-hour SECCHI/HI running difference image showing quiescent solar wind structures. The figure shows that the largest intensity (therefore density) variability corresponds to locations nearest the HCS. In other words, SECCHI/HI can image the HPS directly and detect its intermittent structure. These measurements can identify the sources of the solar wind structures when compared with in-situ abundance SSP measurements from other IHPs.
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SoloHI will have much better sensitivity and spatial resolution than HI. This will allow us to trace the HPS boundaries, their evolution and their relation relative to the HCS in much greater detail than possible with STEREO. When combined with the in-situ observations from the Solar Orbiter, and/or SPP, the SoloHI observations will provide strong constraints on the origin and evolution of the solar wind plasma in the heliosphere. Observations such as those in Figure 2-2 also Figure 2-2. Comparison of SECCHI/HI enable us to obtain measures of the solar wind observation of solar wind structures (grey turbulence directly from the SoloHI images. scale) with the model-derived location of the Because the observed emission is related to the HCS (red surface). The meridional slice number of electrons along the line of sight, shows the model velocity but it is not intensity variations provide a direct measure of discussed here. solar wind density variations which can be compared to Earth-based interplanetary scintillation or Sentinels‟ in-situ measurements. SoloHI can run a specific observing program for this case (wave turbulence programs in Table 2-8). For example, we will use the SoloHI images to predict when Solar Orbiter (or SPP) will cross a solar wind structure of interest (e.g., an HPS boundary or a fast stream interface). For a time interval of about 4 hours before the Solar Orbiter passage, (or around the SPP passage) SoloHI will obtain images over a restricted FOV (about 1.5ºx5º) around the region of interest. A power spectrum of the density fluctuations can then be constructed by using difference images with variable cadences. Such a program is made possible by the programming flexibility offered by the SoloHI electronics (see Section 3.1.7 and Table 3-1).
2.3.2 How do solar explosions produce energetic particle radiation that fills the heliosphere?
Key Science Questions How and where are energetic particles accelerated at the Sun? How are solar energetic particles released from their sources and distributed in space and time? What are the seed populations for energetic particles?
CME-driven shocks play a central role in determining the energetic particle populations in the heliosphere and in driving geospace storms. They are known to accelerate solar energetic particles (SEPs) to high energies (e.g., Reames 1999; Kahler 2001), even GeV energies (Bieber et al. 2004) during the so-called gradual SEP events. Fermi acceleration is the likely acceleration mechanism for quasi-parallel shocks while gradient-drift acceleration operates at quasi- perpendicular shocks (e.g., Lee 2000). The geometry of the shock seems to play a further role in the observed variability of the spectral characteristics and composition of SEPs (Tylka et al. 2005). The shock compression ratio determines the power law index of the SEP spectrum under some simplifying assumptions such as equilibrium conditions. The particles must extract energy
2-5 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 from the shock to reach their observed energies. It appears that the particle kinetic energy might be a fairly significant percentage of the CME kinetic energy (Mewaldt et al. 2005). Many of these shock-related parameters (geometry, compression ratio, speed) are available or can be deduced from in-situ measurements at 1 AU. None, however, is actually measured in the low corona where the highest energy particles originate (≤10 Rsun, Tylka et al. 2005). Moreover, the large scatter in the correlation between CME speeds and SEP peak intensities suggests a complex interplay among the CME speed, the acceleration mechanism(s) and the ambient environment. Recent work has focused on the role of the variations of the environment through which the CME shocks and particles propagate (Gopalswamy et al. 2004; Kahler & Vourlidas 2005). The results indicate that SEP-rich CMEs tend to occur in the aftermath of a preceding CME but they also tend to have much brighter fronts than SEP-poor events. Since bright emission in a coronagraph image implies a large extent along the line of sight, the latter finding suggests that SEP-rich CMEs attained larger longitudinal and latitudinal extents, at lower heights than SEP- poor CMEs. Therefore, the height of formation of the shock, the 3-dimensional extent of the CME, and the existence or not of a preceding event are necessary observations for a better understanding of the generation and propagation of SEPs. The SoloHI instrument will provide these crucial observations. SoloHI will image CMEs and their associated shocks at the coronal heights where the particles originate (≤10 Rsun) with sufficient spatial resolution to resolve the locations of the CME-driven shocks. The optimal period for such observations is during perihelion passages when the SoloHI FOV extends from 5.8 to 42 Rsun with 52 arcsec resolution (AUeq) for 2x2 binned images, the standard observation mode (see example observing program A1.1, Table 2-9). In other words, SoloHI is similar to a LASCO/C3 coronagraph with 2x better spatial resolution. Previous work (Vourlidas et al. 2003) has shown that CME-driven shocks can be easily detected with LASCO/C2. More recently, Ontiveros & Vourlidas (2009) have extracted quantitative measurements of the density profile at the shock front for several CMEs and at heliocentric distances of up to 17 Rsun. In addition, they fit the observed profiles with a simple geometric model of a bow shock, thus estimating the direction of shock propagation. Since the instrument has similar spatial performance but higher sensitivity than the LASCO instruments, SoloHI will readily observe and characterize the evolution of shocks. For example, a cadence of 30 min will allow 6 observations of a 2000 km/s CME in the SoloHI FOV during perihelion. The SoloHI inner field of view will extend below 10 Rsun for all heliocentric distances within 0.5 AU, and therefore will be able to contribute to the SEP analysis for a much larger part of the Solar Orbiter orbit than just at perihelion. For these parts of the orbit, SoloHI will be able to observe shocks and CMEs as they go over at least one of the IHPs. To address the spatial extent of the shocks, the SoloHI observations can be combined with simultaneous observations from the SECCHI coronagraphs and imagers, if the STEREO mission is still operating. The multipoint observations will be used to reconstruct the 3-dimensional structure of the CMEs and their associated shocks. Alternatively, the shocks can be localized with the help of radio observations of type-II bursts from the RAD instrument on Solar Orbiter and the corresponding instruments on SPP and STEREO.
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2.3.3 How do solar transients drive heliospheric variability?
Key Science Questions How do CMEs evolve through the corona and inner heliosphere? How do CMEs contribute to the solar magnetic flux and helicity? How and where do shocks form in the corona and inner heliosphere?
Most observations of CMEs, over the last 30 years, have been made using space-based white- light coronagraphs. The highly successful LASCO coronagraphs have observed over 10,000 CMEs (Yashiro et al. 2004). Many CMEs result from the disruption of a helmet streamer overlying a magnetic neutral line on the Sun (Hundhausen 1993), and they are often observed to have a three-part structure in which a bright leading shell of material surrounds a dark cavity within which a bright core may reside (Illing & Hundhausen 1986). The three parts of the classical CME may be identified with the bright dome, cavity, and quiescent prominence of the helmet streamer. An example of a CME with this three-part structure is shown in Figure 2-3 (Plunkett et al. 2000). Concave-outward striations are observed overlying the erupting prominence, which appear to form a closed circular structure with the bright front of the CME. These features have been interpreted to indicate the presence of a large-scale helical magnetic flux rope in the CME cavity (Vourlidas et al. 2000). The evolution of this flux rope structure close to the Sun is well understood from LASCO observations. Cremades & Bothmer (2004) showed that CMEs expand in a self-similar manner as they propagate outward from the Sun, as models suggest (Low 1982; Gibson & Low 1998), and that they often have a cylindrical geometry, with the axis of symmetry corresponding to the long axis of a large-scale helical magnetic flux rope that originated in the CME source region. Thernisien et al. (2006) applied a forward modeling technique to fit an empirical model of a magnetic flux rope to a sample of 34 CMEs observed by LASCO. By modeling the flux rope as a graduated cylindrical shell of electron density, they were able to reproduce the observed morphology of the CMEs and derive the electron density in the CME front under the assumption of self-similar expansion.
Figure 2-3. Left: A typical 3-part structured CME in LASCO/C2 on June 2, 1998. Right: A similar CME observed by SECCHI/HI1 on July 9, 2007 at a height of ~40 Rsun.
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To understand the propagation and evolution of these events further into the heliosphere, we have until recently relied only on MHD models and in-situ observations of interplanetary CMEs (ICMEs). Imaging observations of CMEs do not provide direct measurements of their plasma and magnetic field properties and have generally been restricted to small elongations from the Sun. On the other hand, quantitative measurements of ICME plasma and magnetic fields are only available from in-situ instruments located far into the heliosphere. This dichotomy restricts our ability to understand the origins and evolution of CMEs. In-situ measurements to date sample only a single trajectory through the large-scale ICME. In the case of simple morphologies, the large-scale structure can be inferred from such measurements, for example by modeling them as force-free magnetic flux ropes (Wu & Lepping, 2006). However, the verification of such analyses is problematic since the events that are best sampled in-situ appear as halos in the coronagraph images and suffer uncertainties in the derivation of their speed and morphology due to severe projection effects. Direct imaging of ICMEs coupled with contemporaneous multi- point in-situ sampling and MHD modeling is the only way to understand how CMEs evolve in the inner heliosphere and to finally be able to connect CMEs to ICMEs.
MHD simulations of CMEs propagating in a structured solar wind show that the motion and local appearance of CMEs are strongly affected by its interaction with the background velocity and density structures (e.g., Odstrcil & Pizzo 1999a, b, c; Riley et al. 2003; Odstrcil et al. 2004; Odstrcil et al. 2005). The initial shape and density distribution of the CME is distorted in all dimensions; it is compressed where the CME is trapped between slow streamer belt and high-speed coronal hole flows, and it is distended where the CME penetrates into the trailing edge of the preceding high-speed stream (Figure 2-4). Some remotely sensed measurements of CMEs at large distances from the Sun have been made, using interplanetary radio scintillation techniques (Manoharan et al. 2001), Helios white-light photometers (Jackson & Leinert 1985; Webb & Jackson 1990) and more recently with the Solar Mass Ejection Imager (SMEI; Webb et al. 2006) and the STEREO/SECCHI HI (Vourlidas et al. 2007). Most CMEs observed by SECCHI show Figure 2-4. Numerical Modeling of the Propagation of the 5/12/97 CME in the significant structural evolution during their propagation in the HI1 and HI2 fields of view. Heliosphere So far, we have found that almost all CMEs lose their internal structure by the time they enter the HI2 field of view (at an elongation of about 20) and typically only an arc-like front or series of fronts remain. Some of these apparent structural changes may be due to instrumental sensitivity effects. Interpretation of the observations is also complicated by the fact that, as CMEs propagate outward from the Sun, the surface of maximum Thomson-scattering efficiency lies along a
2-8 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 sphere centered halfway between the Sun and the observer, with a diameter equal to the distance between the Sun and the observer (the Thomson surface, see Section 2.4.2). Thus a CME that is easily observed in the so-called „plane of the sky‟ near the Sun may have a very different appearance, or may not be visible at all, much further from the Sun, when viewed by an observer in the same location. SoloHI will extend the field of view of the SECCHI HI1 to elongations greater than 40 from the Sun with increased sensitivity. When combined with state of the art MHD simulations to help interpret the observations, SoloHI will provide the best measurements to date of the evolution of CMEs in the inner heliosphere, and of the interaction of CMEs with ambient heliospheric structures. CME speeds observed near the Sun range from less than 100 km/sec to more than 2,500 km/sec (e.g., Yashiro et al. 2004), while the corresponding ICME speeds near 1 AU are typically within 100 – 200 km/sec of the ambient solar wind speed. The aerodynamic drag force due to the interaction of the ICME with the solar wind leads to this equalization of ICME and solar wind speeds (Gosling & Riley 1996; Vrsnak & Gopalswamy 2002; Cargill 2004). Fast halo CMEs are often observed to decelerate in the outer part of the LASCO field of view, when the CME is far from the Sun (Sheeley et al. 1999). Recently, the SECCHI observations have allowed comparisons of speed measurements from remotely sensed and in-situ data (e.g., Rouillard et al. 2009, Wood & Howard 2009) which show good agreement at 1 AU. However, the detailed acceleration profile of individual CMEs in the inner heliosphere remains ambiguous because of uncertainties in the proper identification of CME features due to line-of-sight integration effects. Imaging of CMEs in the inner heliosphere from small distances less than 1 AU are needed. Positively identifying the same CME in SoloHI images and in-situ observations (from Solar Orbiter or SPP or other heliospheric probes) will allow the detailed analysis of CME propagation and evolution in the inner heliosphere.
Various parameters are used to identify the boundaries of ICMEs from in-situ solar wind plasma and magnetic field data, often yielding significantly different results for the same event (Mulligan et al. 1999). These signatures have been used to identify two types of ICMEs that occur in approximately equal proportions in the solar wind: (1) magnetic clouds, whose local magnetic structure is that of a flux rope (Burlaga 1995), and (2) „complex ejecta‟, which are not flux ropes and have disordered magnetic fields (Burlaga et al. 2001). Figure 2-5 shows a schematic of an ICME with a magnetic cloud structure (Zurbuchen & Richardson 2006). The relationship of the three-part structure of CMEs near the Sun to the structure of ICMEs is not well understood. Figure 2-5. Schematic Representation of an Bothmer & Schwenn (1994, 1998) showed that Interplanetary CME the magnetic field structures of solar filaments and their overlying coronal loop arcades are well correlated with the field structure in associated magnetic clouds. Some efforts have been made to compare remotely sensed CME observations
2-9 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 with in-situ ICME observations when the appropriate spacecraft have been in a quadrature configuration (Gopalswamy et al. 2001). Thanks to STEREO, such comparisons can be made at any time (e.g., Kilpua et al. 2009; Möstl et al. 2009; Lynch et al. 2010; Liu et al. 2010, and others).
However, these efforts have been hampered by the large gap in spatial coverage between the coronagraph images and the in-situ plasma and magnetic field data. It is generally assumed that the bright front evolves to become the sheath of compressed solar wind, while the dark cavity corresponds to the magnetic flux rope seen in magnetic clouds. The fate of the dense core seen in coronagraphs is less clear. In-situ observations have, in several cases, revealed the presence of a “plug” of cold, dense plasma trailing the flux rope, which has been interpreted as remnant material from the erupting filament (the bright core seen in coronagraph images) (Gopalswamy et al. 1998; Ho et al. 2000). However, filament plasma (indicated by an enhanced He+/He++ ratio) has not been observed in the vast majority of ICMEs. The question thus arises: what is the fate of the erupting filament as the CME propagates into interplanetary space? The traditional answer is that the filament is so small relative to the CME that spacecraft simply miss it most of the time. However, the bright core often occupies a significant fraction of the CME Figure 2-6. Left panels: SECCHI/HI1 volume in coronagraph observations (Figure observations of the tail disconnection of 2-3). One possibility is that most of the filament Comet Encke. The images are shown in plasma drains down to the surface during the inverted brightness and are histogram- eruption and only a very small fraction escapes equalized to emphasize faint structures. Right with the CME. An alternative explanation is panels: Running difference images. The that the filament plasma is present but has lost vertical streaks are light from bright stars due its expected low-charge-state signature because to the shutterless operation of the of heating (Skoug et al. 1999; Rakowski et al. SECCHI/HI cameras. The faint cloud 2007). SoloHI e will be able to directly image approaching the comet is part of the CME the CME before it passes over the SPP or front. Bepi-Colombo, thus permitting a direct comparison of the remotely sensed and in-situ measurements. An example of the synergy that we can expect between SoloHI and inner heliospheric probe observations was the serendipitous observation by SECCHI HI1 of the interaction between comet Encke and a CME in April 20, 2007 (Figure 2-6, Vourlidas et al. 2007).
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Complex ejecta are most likely formed when two or more CMEs merge in the interplanetary medium to form a single structure. Burlaga et al. (2002) identified three cases in which successive CMEs merged en route from the Sun to the Earth to form complex ejecta. The identity of the individual CMEs in each case was lost by the time they reached 1 AU. The interaction process is non-linear and irreversible; however the details are not well understood. Gopalswamy et al. (2001) have reported cases where two or more CMEs appear to interact along the line of sight in the LASCO coronagraphs, leading to enhanced long-wavelength radio emission. These instances of so-called „CME cannibalism‟ are frequently associated with intense SEP events (Gopalswamy et al. 2004). However, Richardson et al. (2003) pointed out several problems with the interpretation of the observations as being due to direct interaction between CMEs. SoloHI will provide observations of CME interaction over wide range of elongation angles in the inner heliosphere, while in-situ plasma and magnetic field measurements from the Solar Orbiters, SPP, and Bepi-Colombo, arrayed at different radial distances from the Sun, will permit the interaction process, and its effects on SEPs, to be studied in detail. In the absence of some mitigating process, CMEs would continuously add new magnetic flux to the heliosphere, leading to an unchecked buildup of flux that is not observed. Rather, the total unsigned magnetic flux in the heliosphere varies by a factor of about two over the solar cycle (Wang et al. 2000; Smith & Balogh, 2003). The likely explanation is that closed field lines within CMEs partially disconnect from the Sun through reconnection of adjacent loops in a sheared arcade to form a flux rope made up of nested helical magnetic fields, and also undergo interchange reconnection with neighboring open field lines of the opposite polarity (Gosling et al. 1995; Crooker et al. 2002). In-situ observations show that the field in ICMEs opens slowly over many months, and that approximately 50% of the flux in ICMEs remains closed out to 5 AU (Crooker et al. 2004; Riley et al. 2004). If indeed the legs of ICMEs remain connected to the Sun for such an extended period of time, then ICMEs may contribute significantly to the factor of two variation in the heliospheric flux over the solar cycle. With SoloHI, we will be able to follow the morphology of CMEs well into the heliosphere, and we will be able to determine whether some parts of those CMEs in fact remain connected to the Sun over large distances. These observations will provide the necessary context for in-situ measurements from the Sentinels to better understand the field topology within ICMEs, and to establish the contribution of ICMEs to the heliospheric magnetic flux over the solar cycle. 2.3.4 What is the three-dimensional structure of the heliosphere?
Key Science Questions What is the three-dimensional structure and extent of streamers and CMEs? How are variations in the solar wind linked to the Sun at all latitudes? What are the sources and properties of dust in the inner heliosphere, and do Sun-grazing comets contribute to the dust?
Our current knowledge of the sizes of CMEs is limited to projections of their angular and radial extent on the sky plane as viewed along the Sun-Earth line. Although we can get a good estimate of their latitudinal and radial extent from observations of multiple events, we have no direct knowledge of the longitudinal extent of these events. It is, therefore, difficult to investigate the spatial relationship between CMEs and their associated streamers, how CMEs affect the corona, and how quickly the corona recovers after an eruptive event. The answers to these questions are
2-11 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 part of the main objectives of the recently launched STEREO mission, and the first results from the SECCHI coronagraphs (e.g., Thernisien et al. 2009; Mierla et al. 2009; Thernisien & Vourlidas 2010) are very encouraging. However, the STEREO spacecraft constitute only 2 lines of sight through the corona, both of which lie very close to the ecliptic plane providing partially redundant information. Newmark et al. (2003) have shown that the resulting 3D reconstructions have large longitudinal uncertainties which can be improved only with the addition of a third viewpoint, especially from a location away from the ecliptic plane. These problems are exacerbated when we try to measure the 3D structure of streamers because their longitudinal boundaries are more difficult to define. Since streamers are long-lived coronal structures, lasting for several rotations, one can rely on the solar rotation to provide additional viewpoints. These techniques rely mainly on the inversion of pB coronagraph measurements and can achieve a temporal resolution of about 4.7 days (1/6 of the solar rotation) using data from 3 satellites (Frazin & Kalamabadi 2005). They are best suited for coronal reconstruction during periods of minimum activity although new techniques, such as Kalman filtering, seem to be able to capture some of the short-term evolution of coronal structures (Frazin et al. 2005). One should note that the above results are based solely on experience with observations from the ecliptic and at 1 AU from the Sun. Solar Orbiter will provide high latitude views of the extended corona from the surface to 45 elongation, starting in 2020 and possibly extending to 2026. This time period covers the maximum of Solar Cycle 25, when the corona will undergo continuous configuration changes under the influence of frequent CMEs and photospheric flux emergence, and when knowledge of the evolving coronal structure will be vital for the Solar Orbiter objectives. The SoloHI will provide completely new information and constraints for understanding the 3D structure of the corona and CMEs. It will observe the heliosphere from out-of-ecliptic viewpoints, at varying latitudes and heliocentric distances thus providing strong constraints on the longitudinal extent of the structures. The Solar Orbiter-Sun distance plays an important role in the data analysis because it affects the visibility of the Thomson scattered features in the images, as mentioned earlier (Section 2.3.3). In essence, the varying distance acts as a filter emphasizing structures that are progressively nearer the spacecraft as it approaches the Sun. When combined with simultaneous white light observations from a point at 1 AU (e.g., STEREO), the analysis of the emission should provide very strong constraints on the size of coronal features. The idea of the Thomson sphere and the concept of the joint analysis of observations from different heliocentric distances are so new that there exist no simulations or other assessments of their potential on heliospheric reconstructions. SoloHI will be the first instrument to provide high resolution imaging observations of Thomson-scattered emission from out of the ecliptic. The Solar Orbiter mission possesses another important advantage; within two weeks the spacecraft sweeps between its two latitudinal extremes (from south to north) thus enabling tomographic reconstructions of streamers using the SoloHI data alone with the same temporal resolution achievable currently by the LASCO instrument. Besides reconstruction input, SoloHI will provide important direct observations from out of the ecliptic. In particular, SoloHI will be able to observe the interaction of the propagating CMEs with the solar wind and other CMEs from a unique vantage point. Depending on the inclination of the solar dipole during the high-latitude passes, SoloHI will be able to make a direct measurement of the longitudinal extension of the CME, the interaction of the CME with the boundaries of coronal holes and the interaction between CMEs, through a shorter line of sight
2-12 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 than is possible from the ecliptic. The shorter line of sight will minimize uncertainties in these measurements due to the effect of overlapping structures. To get the full benefit of the out-of-ecliptic viewpoint in understanding the structure and dynamics of the corona, the SoloHI observations must be combined with the coronagraph and low corona EUV imager observations. Using the coronal imagers on Solar Orbiter, we can observe the initiation and evolution of Earth-directed CMEs and more easily determine the spatial relation between CMEs and their coronal sources. At or near perihelion when SoloHI will be observing close to the Sun, the combination with the coronagraph and EUV imager will allow us to observe polar plumes and small-scale structures in the polar and equatorial coronal hole regions. Hence we will determine the 3D structure of coronal plumes and perhaps determine their coronal sources.
The question of how the solar wind variations are linked to the Sun at all latitudes encompasses the same detailed issues addressed in Section 2.3.1 but from different, out-of-the-ecliptic viewpoints. The connection between the in-situ solar wind and its solar sources will be analyzed with Solar Orbiter‟s measurements of solar wind speed, energy flux, magnetic field, chemical composition, and ionization state as functions of latitude. However, determining the connection between the in-situ measurements and their coronal sources relies on the coronagraph, EUV Imager, and SoloHI. SoloHI provides the connecting link between the in-situ and solar Figure 2-7. (Top) Solar wind velocity and surface measurements from Solar Orbiter in magnetic field superimposed on composite the inner heliosphere and the inner corona. SOHO EIT & LASCO images contrasting the Ulysses observations have shown only a large- Sun and its wind at the minimum (first orbit) scale, statistical connection between the white and maximum (second orbit) of solar activity. light features and in-situ solar wind Solar Orbiter obtains 2 +/-30° scans per year. measurements (Figure 2-7). The details of this connection remain elusive. With SoloHI, we can fill in the gap between the speed profiles and the coronagraph measurements and trace the coronal sources of the solar wind variation. For example, observations of the solar wind at high latitudes near 0.5 AU may allow the connection between coronal holes and solar wind features to be established. Determining the physics and structure of the wind‟s acceleration requires tracing the outflowing plasma from the solar surface, through the chromosphere and transition region, to the corona using the EUV imager and coronagraph. From there, SoloHI will be able to image the size, shape, and velocity structure of the solar wind inhomogeneities coming from coronal hole boundaries as a function of time, providing critical boundary conditions for coronal and solar wind models. The visible emission at heights above 4 Rsun is dominated by scattering from interplanetary dust, the F-corona. It is a nuisance for coronal studies as it obscures the signal from CMEs and coronal streamers. Accurate removal of the F-corona is essential for the derivation of coronal density structure (e.g., Hayes et al. 2001), but the current F-coronal models are insufficient, as LASCO-C3 observations have shown. It has taken 10 years of C3
2-13 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 observations and the development of several data analysis techniques to achieve an accurate F- corona model. The same techniques are being used for the removal of the F-corona from the SECCHI images. In the case of SoloHI, however, we will be in position to extract quantitative measurements of the F-corona from different places in the ecliptic for the first time since the Helios mission ended in the 1980s but with much higher sensitivity, spatial resolution and spatial coverage compared to the Helios photometers. Moreover, SoloHI will record the first F-corona images from locations outside the ecliptic where little is known about the brightness and spatial distribution of the F-corona emission (Leinert et al. 1998). We do not know, for example, whether long-period comets are an important source of dust in the inner heliosphere. The detection of a significant population of large particles at high ecliptic latitudes would support this idea (Delsemme 1976). The combination of the SoloHI remote-sensing F-corona observations with the in situ measurements from the Solar Orbiter dust instrument will allow the estimation of the size distribution of the dust in the inner heliosphere. In particular, the analysis of the changes in the distribution of particles below a few micrometers will yield information about the effects of solar radiation and plasma environment on the interplanetary dust (Mann et al. 2000). Finally, the SoloHI F-corona observations during the maximum of cycle 25 from high latitudes and within 0.5 AU provide an unprecedented, and probably unique, possibility to investigate whether CMEs interact in any significant way with the interplanetary dust and whether we can use this interaction to probe the CME magnetic fields, as suggested by Ragot & Kahler (2003). 2.3.5 Analysis Techniques Software tools for common analysis tasks that are in use for LASCO and SECCHI may be extended to incorporate SoloHI data. These include image visualization, generation of movies, feature tracking, structure measurement, and combining datasets from multiple remote-sensing and in-situ instruments and spacecraft. Forward fitting of three-dimensional models to heliospheric features such as streamers and CMEs will also be provided in Solarsoft. NRL will work with the Community Coordinated Modeling Center (CCMC) at NASA Goddard Space Flight Center to produce appropriate heliospheric model calculations for comparison with the SoloHI data for each Carrington rotation, as well as for selected events of interest. The results of these model calculations will be made publicly available on the WWW. 2.4 Observational Requirements Table 2-2 provides the Science Requirements Traceability Matrix (SRTM). This table shows the flow down from each of the science objectives discussed in Sections 2.3.1 - 2.3.4 to the instrument performance requirements, and to the observations that will be needed to accomplish the objectives. 2.4.1 SoloHI Signal to Noise The total diffuse sky brightness observed by SoloHI is the sum of contributions from the F- corona or zodiacal light (scattering of sunlight from interplanetary dust) and K-corona (scattering from electrons in the solar wind and CMEs), as well as integrated starlight from stars not resolved by the telescope. The dominant contribution throughout the SoloHI FOV comes from the F-corona. In addition, numerous bright stars will be resolved as individual point sources of light. These stars will be removed from the images using techniques that have been developed and successfully implemented on SECCHI/HI. The fundamental observational requirement is to obtain images with sufficient photometric precision to discriminate the K-corona, and its
2-14 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 fluctuations (e.g., CMEs), from the other contributions to the total signal. The SoloHI instrument has been optimized to observe structures in the solar wind throughout its FOV with a sufficient cadence to measure the properties (speed, direction of propagation, density and internal structure) of CMEs and other transient flows. Figure 2-8 shows a comparison of the 1 detection limit for the SoloHI with the expected signals. The detection limit shown here is for a single pixel and an exposure of 30 minutes at aphelion and 30 seconds at perihelion. The expected contribution of the F-corona was determined from the model of Koutchmy & Lamy (1985). The integrated starlight varies considerably with galactic latitude and longitude; an average contribution over the SoloHI FOV is shown in Figure 2-8. The expected contribution from CMEs is derived from Helios measurements by Jackson et al. (1985), and from SECCHI/HI measurements near solar minimum. The 1 detection limit remains below the expected CME signal over the full field of view. A signal-to-noise ratio (SNR) of 5 per spatial resolution element is required for threshold detection of a simple, known a-priori target on a flat background (Rose 1948; Barrett 1990), and substantially higher photon statistics (SNR>30) are required for more complex or unknown targets. These criteria are easily met over most of the SoloHI field of view with an exposure of 30 minutes at aphelion; near the outer part of the field binning of pixels and/or longer integration times will be used to enhance the desired signal. At perihelion and/or in the inner part of the field, substantially higher cadence using subframes can be achieved without any degradation in resolution.
Figure 2-8. Expected contributions to the SoloHI signal as a function of elongation angle for perihelion (0.22 AU) and aphelion (0.88 AU). The 1- photon noise detection limit per pixel is shown for an exposure of 30 min at aphelion and 30 sec at perihelion. See Figure 3-5 for a revised figure with the correct perihelion distance of 0.28 AU.
2.4.2 Thomson Surface Considerations Solar wind features at progressively large angular distances from the Thomson surface scatter less than features close to the surface. Therefore, the SoloHI measurement sensitivity and its scene coverage is defined by the Thomson surface that varies with the Sun-observer distance. Figure 2-9 shows an estimate of the SoloHI scene coverage for 3 SoloHI orbital positions near its orbit perihelion. There are 3 curves for each SoloHI location that are plotted with the same color and define the distance along the SoloHI line-of-sight, where the integrated scene brightness
2-15 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 reaches 5%, 50%, and 95% of its total brightness integral. 90% of the scene brightness captured by the SoloHI instrument lies between the 2 outermost solid-line curves. The numbers in Figure 2-9 refer to separate locations along the orbit. „2‟ is at the orbit perihelion of 0.28 AU, while „1‟ and „3‟ are both at 0.34 AU. The orbital transit time from position „1‟ to position „3‟ is approximately 17.6 days. As the SoloHI orbital position approaches its perihelion, the SoloHI instrument will primarily measure Thomson-scattered light from within 40 Rsun of Sun center and therefore become a local-space imager.
However, SoloHI can still image both the distant and local environment. SPP will be within the imaging area of SoloHI at certain times during the mission. The duration of the SPP passage across the wide field SoloHI instrument FOV during the Solar Orbiter mission is based on the final SPP launch date and orbit. This analysis demonstrates that SoloHI will easily detect the features passing over SPP while being sensitive to the solar wind flowing towards the Solar Orbiter S/C. 2.5 Science Data & Other Scientific Products 2.5.1 Science Requirements Traceability Matrix (SRTM) Figure 2-10 shows the flow down from the Level Figure 2-9. Sensitivity map of the Thomson 1 program requirements to the Level 3 mission scattering emission for the Solar Orbiter- segment requirements, including the top-level Sun geometries at perihelion. Each number requirements for the SoloHI instrument. The denotes a location in the orbit („2‟ is SoloHI SRTM in Table 2-2 shows the SoloHI perihelion of 0.28 AU, while „1‟ and „3‟ are science measurements and observation both at 0.34 AU). The 3 arches in each color requirements that are flowed down from the each mark the locus of the 5%, 50%, and 95% Solar Orbiter science objective and science brightness integrals along the line of sight. question. Table 2-3 provides a summary of the The extent of the arches marks the SoloHI SoloHI Level 1 science requirements that are FOV. The dotted lines show the direction of shown in Program Level Requirements for the the Parker Spiral for a 300 km/s wind. Solar Orbiter Project Appendix D for the Living With a Star Program.
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Figure 2-10. SoloHI Requirements Flowdown from Level 1 to Level 3
2.5.2 Threshold Science Mission The SoloHI threshold science mission is described by the SoloHI threshold science measurement requirements captured in Program Level Requirements for the Solar Orbiter Project Appendix D for the Living With a Star Program document and the SoloHI Science Requirements Document. Table 2-3 and Table 2-4 capture the SoloHI Level 1 and Level 3 threshold science measurement requirement respectively. The SoloHI instrument performance to satisfy the threshold science mission is nearly identical to the SECCHI HI1 instrument performance for the STEREO mission, as summarized in Table 2-1. Although the telescope performance is considerably reduced, the minimum mission SoloHI will still obtain breakthrough observations of the inner heliosphere. Each of the reductions has a different scientific impact. Raising the inner field limit from 5.5° to 6.0° increases the gap between the SoloHI and the Solar Orbiter coronagraph observations and subsequently complicates the task of tracking structures from the low corona into the solar wind. The threshold mission field of view is the same as the STEREO HI1. This reduced field restricts the range of distances over which solar wind features can be tracked. The reduced spatial sampling of the
2-17 Version 2.0 - 31 December 201 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 image increases the spatial sampling to be roughly equivalent to the geometric image quality of the telescope with only a modest loss of spatial resolution. The SoloHI minimum mission meets all of the CME-related measurement objectives (Section 2.3.3) and two thirds of the solar wind and SEP-related objectives (Sections 2.3.1 and 2.3.2). However, SoloHI can meet all of its out- of-the-ecliptic objectives (Section 2.3.4) with its minimum mission. Table 2-1. SoloHI Threshold Mission Measurement Requirements Comparison to STEREO HI
Parameter Nominal Threshold STEREO HI1 Units Inner Limit 5.5 6.0 4 Degrees from Sun Center Field of View 40 20 20 Degrees (Full Width)
Spatial Sampling 50 61 72 Arc Sec (1 AUeq) 2.5.3 Descope Options Table 2-5 identifies the SoloHI descope options that will violate the SoloHI baseline science measurement requirements, but still satisfy the SoloHI threshold science measurement requirements. This table describes the science impact and programmatic impact, if any of these descope options are exercised by the SoloHI Project Management. In addition, the schedule deadline for the trigger date to exercise each descope option is defined. These descope options will reduce instrument resources, including mass and envelope, and will reduce technical risk in satisfying the top-level instrument photometric accuracy and spatial resolution requirements. The only descope option that offers a modest cost savings to the SoloHI baseline budget is Descope Option 4 that degrades the image spatial resolution requirement.
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Table 2-2. SoloHI Science Requirements Traceability Matrix (1 of 4)
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Table 2-2. SoloHI Science Requirements Traceability Matrix (2 of 4)
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Table 2-2. SoloHI Science Requirements Traceability Matrix (3 of 4)
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Table 2-2. SoloHI Science Requirements Traceability Matrix (4 of 4)
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Table 2-3. SoloHI Level 1 Science Measurement Requirements for the Baseline and Threshold Science Mission
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Table 2-4. SoloHI Level 3 Science Measurement Requirements for the Baseline and Threshold Science Mission (1 of 2)
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Table 2-4. SoloHI Level 3 Science Measurement Requirements for the Baseline and Threshold Science Mission (2 of 2)
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Table 2-5. SoloHI Program Descope Options
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2.5.4 Baseline Observing Programs The SoloHI baseline observing program will be designed to satisfy the observing program requirements in the SoloHI Baseline science measurement requirements in Table 2-4 and to satisfy the SoloHI telemetry allocation of 20.5 kbps over 30 observing days of each Solar Orbiter orbit. Many of the baseline science measurement requirements (including radial scene coverage, photometric accuracy, image cadence, and science observation days for the orbit and mission) are dependent on the instrument distance from the Sun and the solar latitude for the SoloHI instrument. For this reason, the observing program for each orbit will be divided into five regions based on the instrument distance from the Sun: • Region 1 (Perihelion): 0.28 to 0.29 AU • Region 2 (Near Perihelion): 0.29 to 0.36 AU • Region 3 (Far Perihelion): 0.36 to 0.42 AU • Region 4 (Southern Out-of-Ecliptic): 0.42 to 0.50 AU • Region 5 (Northern Out-of-Ecliptic): 0.50 to 0.70 AU The observing programs for all five regions can be used for any instrument distance from the Sun from 0.28 AU to 0.70 AU. The radial scene coverage for each of these regions at the SoloHI required inner and outer FOV is defined in terms of elongation (deg) and solar radii on the Thompson surface for the given line of sight in Table 2-6. The radial coverage for each of the different SoloHI images (full frame, inner FOV subframe, and radial swath subframe) is included in this table. (The Thompson surface for a given instrument line of sight is defined where the Thomson scattering angle is 90°.) The SoloHI images are required to be captured from the Solar Orbiter perihelion of 0.28 AU to 0.70 AU. The SoloHI full frame images will capture the radial scene from 5.8 to 76.4 Rsun at the Thomson surface for instrument distances from the Sun of 0.28 to 0.70 AU. The SoloHI inner FOV subframe images will capture the radial scene at the Thomson surface from:
• 6.1 to 10.3 Rsun, for the inner FOV subframe image, centered at 7.0 Rsun at perihelion
• 14.1 to 20.5 Rsun, for the inner FOV subframe image, centered at 15.0 Rsun at perihelion
• 19.1 to 27.0 Rsun, for the inner FOV subframe image, centered at 20.0 Rsun at perihelion for instrument distances from the Sun of 0.28 to 0.36 AU. The SoloHI radial swath subframe image will capture the radial scene from 5.8 to 54.6 Rsun at the Thomson surface for instrument distances from the Sun of 0.28 to 0.50 AU. Figure 2-11 shows the baseline orbit for the Solar Orbiter mission for a launch date of January 1, 2017 that was described in the Solar Orbiter: Consolidated Report on Mission Analysis (SOL- ESC-RP-05500). The Solar Orbiter is divided into five mission phases, including Launch and Early Orbit Phase (LEOP), Observatory Commissioning Phase (or Checkout and Verification Phase (CVP)), Cruise Phase, the Baseline Science mission and the Extended Science Mission. These are described in more detail in Section 2.6.4. Figure 2-11 has divided the Solar Orbiter orbit into its Early Mission/Cruise (includes LEOP and CVP), Baseline Science mission and Extended Science Mission phases. There are 8 orbits (with 8 perihelion passages) in the nominal
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Baseline Science Mission and 4.5 orbits (with 5 perihelion passages) in the nominal Extended Science Mission. If these 12.5 orbits are plotted as a function of the instrument distance from the sun and the solar latitude, as shown in Figure 2-12, it is obvious that there are only 4 unique orbits in the combined Solar Orbiter baseline and extended science mission. Specifically, there are 2 unique orbits for the Baseline science mission (Orbits I and II) and 2 unique orbits for the Extended science mission (Orbits III and IV). The orbital period, the spacecraft distance from the Sun and the extreme Southern and Northern Solar Latitude with their corresponding spacecraft distance is listed for each individual orbit in Table 2-7. The SoloHI baseline observing program will be defined to repeat for each occurrence of the same unique orbit. Therefore, a SoloHI baseline observing program will be defined for Orbit I and will be executed for orbits 1 through 4 of the Baseline Science Mission. A second SoloHI baseline observing program will be defined for Orbit II that will be executed for orbits 5 through 8 of the Baseline Science Mission. Table 2-8describes the SoloHI baseline observing program for Orbit I, while Table 2-9 describes the SoloHI baseline observing program for Orbit II. The SoloHI data volume allocation is 53.14 Gbits over 30 days of observations for each orbit, or an average data rate of 20.5 kbps over 30 days. The SoloHI baseline observing program, including the science data, housekeeping data, and 5.18% CCSDS packet overhead, is constrained to fit within this data volume allocation, as shown in Table 2-10 for Orbit I and in Table 2-11 for Orbit II. These tables show that the SoloHI average data rate for a shorter time period can far exceed, or fall well below, the average data rate of 20.5 kbps over the full 30 days for each orbit. For example, the average data rate over the 4-day Perihelion observing period in the Orbit I Baseline observing program is 31.6 kbps, while the average data rate during the 2-day Southern observing period in Orbit I is only 15.2 kbps. 2.6 Mission Operations 2.6.1 Overview Our mission operations approach is based on our experience with successful operations of solar imaging instruments on the SOHO and STEREO missions. The SoloHI science objectives will be achieved with automated ground and flight systems that offer flexibility to the science planner, while ensuring instrument health and safety. The SoloHI will be designed to operate primarily in an encounter observing mode for the three 10-day observing periods during each orbit. We will baseline an observing program that will meet the baseline science measurement requirements for each orbit during the science mission. Observing sequences tailored to specific science questions will be interspersed with synoptic observations as dictated by mission phase and available resources such as onboard recorder storage space and downlink capability. For example, different science questions will be given priority near perihelion and at the largest latitude passages (see discussions in Section 2.3). The observing program will change as the mission progresses, e.g., as latitude increases, and/or for different parts of the orbit to meet different science objectives. 2.6.2 Science Operations Figure 2-13 identifies the interfaces and concept of operations between the SoloHI Instrument Science Planning Center (ISPC) and Instrument Science Data Center (ISDC) at NRL and the Solar Orbiter Science Ground Segment (SGS). All receipt of telemetry from and commanding of the SoloHI instrument is via the Solar Orbiter Science Operations Center (SSOC) at the
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European Space Astronomy Center (ESA/ESAC) near Madrid. The Solar Orbiter Mission Operations Center (SMOC) and the spacecraft communications infrastructure are provided by the European Space Operations Center (ESA/ESOC) in Darmstadt. We support their efforts by providing a User‟s Manual (including packet definitions), an On-Orbit Operations Handbook, and an On-Orbit Contingency Handbook that documents operational modes and constraints. The SoloHI team supports Solar Orbiter spacecraft and science operations meetings and functions as required. The ISPC provides instrument commands to the SSOC to be coordinated with other instruments‟ commands for submission to the SMOC as consolidated payload operations requests. The SMOC in turn processes and merges the submitted requests into a timeline to be uplinked to the spacecraft. The ISDC receives and archives all instrument telemetry from the SSOC and provides data products to the science community as well as to the Solar Orbiter Data Archive at ESAC. 2.6.3 Planning, Command Generation, and Telemetry Handling NRL will provide all operations personnel, software, and tools for SoloHI science planning, instrument commanding, and telemetry processing. The ground system to be used during flight operations will be identical to that used during I&T before launch. Members of the operations team will be responsible for adapting existing tools from SECCHI and SOHO and also operate the instrument during I&T, ensuring a smooth transition to post-launch operations.
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Table 2-6. Radial Scene Coverage for SoloHI Images for Each Orbital Region Boundary
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Figure 2-11. Spacecraft Distance from Sun During Solar Orbiter Mission
Figure 2-12. Spacecraft Distance from Sun and Solar Latitude During Solar Orbiter Mission
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Table 2-7. Mission Phases and Individual Orbit Timeline During Baseline Solar Orbiter Mission with a Jan 1, 2017 Launch Date
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Table 2-8. SoloHI Baseline Observing Program for Orbit I of Solar Orbiter Mission
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Table 2-9. SoloHI Baseline Observing Program for Orbit II of Solar Orbiter Mission
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Table 2-10. SoloHI Data Volume Estimate for Orbit I Baseline Observing Program
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Table 2-11. SoloHI Data Volume Estimate for Orbit II Baseline Observing Program
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Figure 2-13. SoloHI Concept of Operations The SoloHI ISPC and ISDC will be co-located at NRL. The early operations will be conducted from the SMOC as necessary. A small team of dedicated operators will be responsible for day- to-day operations, under the direction of a Lead Operations Scientist. The team executes the science observing plan, ensures instrument health, and interfaces with the SSOC/SMOC. The SoloHI operations team will participate in the Solar Orbiter Science Operations Working Group (SOWG) for observation and operations coordination with the spacecraft and other instruments. The operators will perform the following tasks: command generation, data processing, data distribution, data storage/archival, and instrument performance evaluation. Command generation and telemetry decommutation will be done using the Integrated Test and Operations System (ITOS). ITOS is an existing, supported, software package designed for that purpose. The SoloHI team has considerable experience with ITOS, having used it with SECCHI/STEREO. ITOS supports translation to/from packets utilizing telemetry and command databases, STOL scripting, and telemetry displays. This same software will be used for software development, ground testing, flight commissioning, and flight operations. The build strategy and software quality tools follow that of the FSW (Section 3.2.1). Schedule planning, data reduction and pipeline processing use the Interactive Data Language (IDL). The ISDC is configured for autonomous data operations (similar to SECCHI) with automatic processing of received telemetry. The command and telemetry (C&T) software monitors housekeeping parameters for status and health, and the operations staff receives immediate automatic notification via text message/e-mail when an out-of-limit condition occurs. Operators have remote access to telemetry displays via an existing Web interface. Tools for plotting trends over time are available. Color coding on telemetry displays provides visual alerts for out-of-limit conditions. The ISPC develops operational instrument commands and scripts. Real-time command upload sequence generation may utilize graphical user interfaces (GUIs). However, during normal operations when real-time commanding is not possible due to light-travel time delays, commands
2-37 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 will be software-generated based on script tables. Critical commands require the operator to confirm intent, and ensure they are not sent inadvertently. (No SoloHI command may cause a loss of mission or loss of the instrument.) We expect infrequent command loads to change the schedule as a function of orbital phase. Schedule validation is performed in software before it is formatted for transfer to the SSOC. Following command transmission, the operator, with the help of software, checks telemetry to verify that the load was received on-board by the instrument and the expected response occurred. All computers and networks are protected by system and network firewalls, monitoring, and regular backups. There are multiple levels of security within the ground system with the most protected area being command uplinks. All networks are isolated from the NRL campus by routers with access limited to the planning, engineering support, and ISPC team. 2.6.4 Mission Phases Table 2-12 shows the nominal operational strategy for SoloHI at each of the mission phases identified in the ESA Solar Orbiter EID-A, Issue 2, Revision 5. During the launch, early operations and transfer phase, the SoloHI door will remain closed to permit outgassing of the instrument and the surrounding spacecraft volume and to keep the instrument temperatures within survival limits with minimal power consumption. Only minimal operations are anticipated during this phase, consisting of initial turn-on of electronics and camera subsystems and flight software checkout. Door-closed commissioning operations will be conducted during selected periods in the first 90 days following LEOPS and during the cruise phase, when available telemetry, spacecraft power and other resources permit (for example, we anticipate that instrument operations will not always be possible during cruise phase). The SoloHI door will open 1-7 months prior to start of the Nominal Science Phase. After open-door commissioning activities, routine observations designed to meet the science objectives discussed in Section 2.3 will be conducted. Table 2-12. Nominal Operational Strategy for the SoloHI at Each of the Mission Phases
Mission Phase Duration SoloHI Operations Launch and Early Launch to L+7 days Minimal or none Operations (LEOP) Checkout and Verification Activation and commissioning L+7 to L+90 days Phase (CVP) as allowed Commissioning and science Cruise Phase L+90 days to Venus GAM2 observations as allowed VGA3 to End of Nominal Synoptic and tailored science Nominal Science Phase Mission (ENM) observations Synoptic and tailored science Extended Science Phase ENM to Venus Encounter #6 observations The standard mode of operations will be to take short exposures (up to 30 seconds) and sum up to N individual exposures on board, with cosmic ray scrubbing carried out on each image. The techniques that will be used for on-board image summing and scrubbing have been developed
2-38 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 and validated on SECCHI. The SoloHI will be operated during three separate 10-day periods throughout each 150 day orbit (30 days per orbit). Data will be stored on the onboard recorder for transmission to the ground during the expected daily four-hour contact period. Highly compressed subsets of the science data can be provided as a space weather beacon for near real- time transmission to Earth when spacecraft operations allow. Nominally the same set of synoptic observations will be taken during each orbit with similar characteristics. In addition to the synoptic observations, a number of critical periods have been identified during which we expect to increase the frequency and amount of SoloHI data above that contained in the baseline synoptic program. High precision images are used to determine the image background. These background images have been very useful to enhance both SOHO and SECCHI images. For SoloHI the background consists of a complex moving stellar background superimposed on the slowly varying (due to the varying radial distance from the Sun) zodiacal light/F-corona. The resulting background is subtracted from individual frames to reveal the Thomson scattered electron signature in CMEs and the solar wind. Observations near perihelion will focus on science objectives that benefit from higher spatial or temporal resolution (e.g., Science Questions 2.1.3 and 2.2.3 in Table 2-2). At this point in the orbit, the S/C will be nearly co-rotating with the Sun, so it will be possible to track individual structures in the SoloHI FOV for up to two weeks. Full resolution images will be obtained, and the field of view will be limited if necessary by transmitting only a portion of each image to the onboard recorder. Special configurations of interest will occur at quadrature, and conjunction with one or more IHPs such as SPP or Bepi-Colombo. (Quadrature occurs when the probe lies on or near the Thomson surface.) Additional observations may be scheduled during these periods and during other periods of joint operations between multiple missions. The periods near maximum northern and southern heliolatitudes will be of particular interest for out-of-ecliptic observations of heliospheric structures. This will be particularly significant during the later part of the nominal science phase, and during the extended science phase, when the orbit inclination increases. 2.7 Data Reduction and Analysis 2.7.1 Overview The ISDC at NRL processes SoloHI telemetry and distributes data products to its Co- Investigators and the Solar Orbiter Data Archive. The basic science product is the 2D intensity distribution, from which a 2D distribution of electron density is computed by assuming that the electrons are in a single volume element on the Thomson surface. The 3D electron distribution can be determined using software that we will make available in the Solarsoft library. 2.7.2 Data Flow The SSOC will receive the instrument telemetry packets from the SMOC and forward them to the SoloHI ISDC at NRL. The housekeeping telemetry will be archived as telemetry monitors in a database from which they are available on-site or over the web. The science telemetry will be processed into Level 1 image files (FITS) with added metadata containing sensor settings, etc. and distributed as described below. Level-1 indicates raw data numbers (DN) without calibration
2-39 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 applied (as defined in the Solar Orbiter EID-A, Issue 2, Revision 5, p127). The metadata keywords will follow the FITS standard augmented with keywords defined in the SoloHI FITS Keyword Definition document. All instrument related information necessary for adequate interpretation and calibration will be distributed through the ISDC website. 2.7.3 Scientific Data Reduction The SoloHI team members are experienced in the reduction and interpretation of white light coronagraph and heliospheric imager data. Science telemetry is received from the SSOC as raw packets in a file or via a socket stream. C&T software strips the packet headers and writes raw data files with accompanying header files, replicating what was on board before being packetized. The data files are read in with IDL procedures which reformat them and write them as Level 1 FITS files with available header information. The Level 1 data can then be displayed and/or calibrated (Level 2+) with the data analysis software package based on IDL and distributed as part of Solarsoft (see Section 2.3.5, Analysis Techniques). Modified or existing routines will be used to correct for missing data blocks, background counts, flat-fielding, and stray light. In order to accommodate the expected range of coronal brightness across the field of view, the output signal from the camera will be digitized to 14 bits per pixel. Images will be summed on board to achieve the desired SNR for the particular science objective for the distance of the spacecraft from the sun. The images will be summed up to 24 bits per pixel. A single image (4Kx4K) with no compression thus takes about 402 Mbits. Lossless compression on the order of a factor of 2.5 can be achieved using the Rice algorithm, which has previously been used on both LASCO and SECCHI. Further reduction in the telemetry rate will be achieved by binning pixels on board for most images (2x2 or 4x4 pixel binning to reduce the image size to 2Kx2K or 1Kx1K pixels) and/or by using a lossy compression scheme such as H-compress where appropriate. Based on these considerations, we estimate a total telemetry volume of about 53 Gbits per orbit for SoloHI data, including CCSDS packet overhead and housekeeping data. 2.7.4 Distribution The SoloHI data policy has completely open access to all data, including the calibration data and all procedures to calibrate and perform high-level processing of the data. SoloHI images will be available as soon as the routine processing steps have been completed (usually within 3 hours of receiving telemetry). After the final images are available, the Level 1 FITS files will be delivered electronically from NRL to the Solar Orbiter Data Archive. These data will be available within the required 3 month timeframe with a goal of 1 – 2 weeks. NRL will also maintain a web interface to a database of all science and housekeeping data that will permit users to search for data corresponding to time periods or events of interest using selected values from the image header, as well as to perform trend analysis of instrument housekeeping parameters such as temperatures and voltages. Validated science data will be distributed directly from NRL to requesters based on the results of a database query. Requests for larger amounts of data will be handled through the Solar Orbiter Data Archive or the Virtual Solar Observatory (VSO). There will be two versions of SoloHI processed science data: quicklook data produced immediately upon receipt of all necessary telemetry from the spacecraft, and final data incorporating any telemetry packets that may be missing or corrupted in the initial telemetry and that are later recovered. Quicklook data will be used for mission operations planning purposes,
2-40 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 and will be available immediately for scientific analysis. Final data will replace the quicklook data when they are available, and will be suitable for archiving and distribution. Both quicklook and final data will be processed in the same way and will have the same file formats. The quicklook FITS file will be differentiated from the final data product by the completeness of the header. NRL is responsible for delivery of the final Level 1 FITS files, as well as all calibration data, to the Solar Orbiter Data Archive. 2.7.5 Calibration The radiometric calibration of the data will be performed using the pre-flight laboratory calibration data and calibration updates using observations of an ensemble of stable stars as used for SOHO/LASCO and STEREO/SECCHI. The calibration team monitors the detector telemetry and the images and provides periodic updates to the science calibration routines. IDL procedures will be provided in the Solarsoft library to convert the Level 1 FITS image files into higher-level calibrated data products. These procedures will permit the user to perform standard corrections such as removal of geometric distortion, vignetting and stray light, and photometric calibration, on the fly for the data of interest. All calibration data necessary for these corrections will be included as part of the Solarsoft distribution which is publicly available at http://sohowww.nascom.nasa.gov. This approach has been used successfully for both LASCO and SECCHI, and ensures that the user has access to the most up-to-date calibrations while avoiding repeated processing and redistribution of large amounts of data. 2.7.6 Data Analysis The project has budgeted adequate resources for mission operations, data processing, analysis and archiving. The resources allocated are based on our recent experience in the development of the STEREO/SECCHI analysis and archive process as well as our continuing analysis and work in support of the SOHO/LASCO data archive. See Section 2.3.5, Analysis Techniques, for a discussion of analysis tools and software. 2.7.7 Database Tools The SoloHI project will use an open source database program such as MYSQL, which is currently being used to manage the housekeeping and image header information on both SECCHI and LASCO. Web-based query tools as well as IDL based tools that are currently available can be easily modified to accept the SoloHI data. The web-based tool enables searches of the image header database with the ability to select FITS files for download using FTP to the user‟s computer. The table structures will be similar to the existing tables. For example, the existing IDL tools include the ability to extract any parameter(s) of interest and to generate plots against time or to correlate one parameter against another. 2.7.8 Computing Facilities The 30-day per orbit operation will result in an uncompressed science data volume of about 90 Gbytes for the eight orbits of the Nominal Science Mission, which is less than 0.3% of our current capacity (as of 12/2010). We will dedicate a single redundant computer system to the processing and storage of the SoloHI data and its products. 2.8 Data Archive and Data Products A complete archive (data, metadata, planning documents, analysis software, etc.) will be maintained at NRL at least during the full mission lifetime. The final instrument calibration and a
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“best and final” calibrated data set (Level 2, as defined in the Solar Orbiter EID-A, Issue 2, Revision 5, p127) from the entire mission will be delivered at the end of Phase E or the extended mission. Prior to that, all image data may be accessed automatically via the Internet either from NRL or from the Solar Orbiter Data Archive. 2.8.1 Data Products Science data from SoloHI will be stored and distributed as uncompressed, uncalibrated Flexible Image Transport System (FITS) files. This product is referred to as Level 1 data. One FITS file will be generated for each image in the spacecraft telemetry stream. The FITS file headers will include keywords to indicate instrument orbit and attitude information, all instrument settings associated with the image, information on all onboard and ground processing steps, image statistics, and any other ancillary information necessary to interpret the image data. Information about the individual exposures used to compute a single image from a sequence will be stored as an ASCII table extension to the standard FITS header. Housekeeping data will be extracted from the raw spacecraft telemetry and will be stored in separate files. The routine processing flow will produce the Level 1 FITS images in addition to several other types of products. These additional products will include browse images and movies for immediate posting on the Internet, higher resolution movies for research, Carrington and synoptic maps of heliospheric brightness at selected elongation angles throughout the field of view, and ancillary data (housekeeping tables and plots, attitude and orbit files). Lists of various events of interest, such as CMEs and comets, will be generated. Table 2-13 gives an overview of SoloHI data products and their formats. 2.9 Science Team Table 2-14 lists the co-investigators and other members of the science team, both funded and unfunded for the SoloHI project. Almost all of the science team listed in these tables will be involved in some aspect of the data analysis tasks, and so this role has not been explicitly identified, except if that is their only role. The PI is also committed to participate in Science Working Team meetings as appropriate. Table 2-13. SoloHI Data Products and Formats
Level Source Description 0 SSOC CCSDS data packets FITS files with uncompressed images. Values are in raw counts 1 SoloHI ISOC (DN) User workstation FITS files with calibrations applied „on the fly‟. Values are in 2 with Solarsoft units of brightness. Data products are the result of combining two or more images User workstation (movies, Carrington maps, etc.) or Derived quantities (electron 3 with Solarsoft or densities, CME masses, etc.). May or may not be calibrated in ISOC website physical units.
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Table 2-14. A World-Class Science Team with Expertise in Instrument Development and Solar Science
Name/Institution/ Role [Funding Capabilities and Experience Justification Source] Instrument Management and Implementation PI LASCO/SOHO and STEREO/SECCHI, Principal R. Howard, PI, NRL development of coronagraphs, CCD cameras, Investigator [NASA] flight software, analysis and interpretation of See Table 4-2 for coronagraph CME and solar wind observations responsibilities Solar radiophysics, CME energetics, coronagraph Deputy PI A. Vourlidas, Deputy development, Project Scientist VAULT and See Table 4-2 for PI, NRL [NASA] VERIS sounding rocket payloads, Project responsibilities Scientist STEREO/SECCHI S. Plunkett, Project Analysis of solar corona, CMEs, space weather, Project Manager Manager & Mission experience in development of space See Table 4-2 for Ops Lead, NRL instrumentation, Co-I and Operations Scientist responsibilities [NASA] STEREO/SECCHI, Co-I LASCO/SOHO PI (VAULT, VERIS) and Co-I (HRTS) sounding rocket payloads, U.S. Project Scientist Instrument C. Korendyke, Hinode/EIS, Co-I STEREO/SECCHI, Scientist Instrument Scientist, LASCO/SOHO instruments, instrument See Table 4-2 for NRL [NASA] laboratory preparation, calibration, and responsibilities operations Instrument Scientist for RAIDS and the Middle J. Morrill, CoI, High Resolution Spectrograph Investigation See Table 4-2 for NRL [NASA] (MAHRSI), Calibration and scientific analysis of responsibilities coronal structures MHD Modeling Scientific analysis of solar and heliospheric data P. Liewer, Lead JPL (Phase E only) and comparison with models, Co-I CoI, JPL [NASA] See Table 4-2 for STEREO/SECCHI responsibilities Solar Wind M. Velli, CoI, JPL Non-linear dynamics and turbulence in space Modeling (Phase [NASA] plasmas, solar activity E only) Analysis and interpretation of CME and solar wind observations, affiliated scientist Solar Wind N. Sheeley, CoI, SWICS/Ulysses, Co-I XUV/Skylab, Modeling (Phase NRL [NASA] Solwind/P78-1, LASCO/SOHO, E only) STEREO/SECCHI
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Name/Institution/ Role [Funding Capabilities and Experience Justification Source] SECCHI/COR2 Instrument Scientist, Optics D. Socker, CoI NRL SECCHI/HI Inst. Designer Consultant R. Harrison, RAL SECCHI Co-I, SECCHI/HI Instrument Lead Data Analysis [UK] P. Rochus, CSL Analysis & SECCHI Co-I, SECCHI/HI Test and Calibration [Belgium] Calibration J-P Halain, CSL SECCHI/HI Test and Calibration Calibration [Belgium] F. Auchere, IAS SOHO/EIT PI, SECCHI CoI Data Analysis [France] P. Lamy, LAM Data Analysis & SOHO/LASCO Co-I, C2 Lead, SECCHI Co-I [France] Calibration V. Bothmer, Univ. of Data Analysis & SECCHI Co-I, Sesame Lead Gottingen [Germany] EPO Collaborators Detector J. D. Moses, NRL SECCHI/Project Scientist; SOHO/EIT Scientist Consultant G. Stenborg, George Development of analysis techniques for coronal Analysis Support Mason University and heliospheric data, Operations Scientist Data archive (GMU). [NRL] LASCO/SOHO support Development of coronagraphs, instrument Calibration and A. Thernisien, GMU laboratory preparation, calibration, development Modeling [NRL] of modeling techniques for coronal and Data archive heliospheric data support Co-I on SECCHI, MER Image Processing, Visualization and E. DeJong, JPL Stereographic Imaging. analysis support
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3. TECHNICAL APPROACH 3.1 Design Approach
The SoloHI instrument concept is based on the successful Heliospheric Imager (HI) currently in operation on NASA‟s STEREO Mission. SoloHI builds on this successful design and provides the opportunity for enhanced observational capabilities in a simplified design. A physical block diagram is given in Figure 3-1. The SoloHI instrument consists of two boxes – the SoloHI Instrument Module (SIM) and the SoloHI Control Electronics (SCE). An optomechanical layout of the SIM is given in Figure 3-2. Instrument design characteristics are given in Table 3-1. 3.1.1 SoloHI Instrument
The SoloHI is a straightforward implementation of the HI concept. The optical train of the instrument is monolithic and incorporates no moving parts. A simple one-shot door based on the same design principles utilized in STEREO/SECCHI provides the necessary protection against contamination during launch, ascent, and AIV activities. The detector is a customized active pixel sensor (APS) provided by Sarnoff Corporation. The detector format is a 4Kx4K with 10 micron pixel pitch. The device architecture is relatively simple and robust. The Sarnoff family of Minimal devices has survived radiation doses up to 2 Mrad with reasonable performance. The progressive scan and reset/row control capability of the APS eliminates the requirement for a shutter mechanism. The control electronics are built around an FPGA based implementation of the ESA-RTC and controls the entire package, acquires the images from the APS camera, performs on-board summing and pixel binning to increase SNR, and transmits CCSDS packets to the S/C. A SpaceWire interface provides the instrument and spacecraft digital communication link. The instrument performance, requirements and budgets are presented in Section 3.1. The telescope and baffle designs are discussed in Section 3.1.4. The structure is discussed in Section 3.1.5. The APS is described in Section 3.1.10 with further details given in the technology development plan, the Sarnoff design reports and the flight detector specification (available on request). The SCE is described in Section 3.1.7 and the flight SW in Section 3.2. During Phase A, we have refined the instrument design. The design is relatively advanced. The text and figures herein represent a snapshot of this evolution. Modest changes and updates continue to be incorporated into the design as the team progresses through Phase A and finishes its preparations for Phase B. 3.1.1.1 SoloHI Design Modifications since the December 2010 CSR
The SoloHI design was significantly updated since December 2010, In February 2011, ESA and NRL reached an agreement to: • Transfer the SIM bracket design, fabrication and test responsibilities from Astrium to NRL, • Move the SoloHI Control Electronics from the spacecraft interior to the Spacecraft exterior on the +Yopt spacecraft panel under the SIM • Increase the SoloHI mass allocation from 14.5 kg to 15.5 kg • Increase the SoloHI survival heater average power allocation from 7.0 to 8.4 W
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The SoloHI design also was subjected to a SoloHI rescope exercise and delivered a white paper on various SoloHI options that would result in cost savings for Phases B through D. The GSFC Solar Orbiter Collaboration Office reviewed these options and formally directed the SoloHI team on June 22, 2011 to adopt the following changes to the instrument design: • Consolidate the Processor card from the former SoloHI Control Electronics (SCE) into the Camera Interface Electronics (CIE). This will eliminate the Cameralink interface between the SCE and CIE. • Eliminate the flight calibration LEDs that were controlled by the SCE. These changes resulted in the renaming of the Camera Interface Electronics (CIE) to the SoloHI Control Electronics (SCE) and of the SoloHI Control Electronics box (SCE) to the SoloHI Power Supply (SPS). The mass of the SIM brackets, the SIM instrument mounts, the SIM structure and the SIM FPA enclosure increased to support the larger SCE mass that is supported by the SIM FPA aft panel. The SoloHI instrument design has also matured in the last year. For example, the operational amplifiers and A/D converters on the DRB were not thermally isolated from the detector in the FPA in the December 2010 SoloHI CSR and were predicted to have temperatures outside of their operational and survival temperature limits. The SoloHI FPA was redesigned in the last year, such that all FPA electronics and components remain within their operational and survival temperature limits with the required margin. In addition, the optical coatings for the SoloHI interior baffles have been changed in local regions to achieve the required straylight rejection at the SoloHI detector.
Figure 3-1. Physical block diagram. Diagram shows the SoloHI components. The SoloHI consists of the SoloHI Instrument Module (SIM) and the SoloHI Power Supply Box (SPS).
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Figure 3-2. SoloHI instrument module optomechanical layout. The SIM design is compact with straightforward optomechanical interfaces. Table 3-1. SoloHI Nominal Design Characteristics Table Telescope Wide angle lens, f=59 mm, aperture=2.418 cm2, aperture stop placed in front of lens, nominal bandpass 500-700 nm. Plate Scale Nominal 35 arcseconds/pixel FOV 40 x 40 square, optimal imaging performance over 24 half angle cone. Image Quality <12 microns rms spot size predicted from ray trace. Detector Active Pixel Sensor, 10 micron pitch, 4096x4096 pixels. Door mechanism One shot door, High Output Paraffin Actuator. Structure Carbon Fiber Reinforced Plastic structure, FPA housing titanium. Low distortion titanium mounts. Baffle Design, Stray Front heat shield edge, forward baffle and diffraction light trap designed to Light Rejection reject incoming solar radiation. Interior baffles and aperture enclosures designed to reject scattered solar radiation from S/C solar arrays and RPW antenna. Peripheral baffles designed to reject radiation scattered from other parts of the S/C. Pointing Instrument axes aligned to S/C to within tolerance, F1 and heat shield leading edge placement error aligned to within tolerance. Tolerances given in EID B. Baffle design accommodates 1.0 excursion from sun center at perihelion. SoloHI will survive excursions of 15 for <20 sec at perihelion and longer excursions at aphelion. SoloHI is not designed to operate at 15 offpoints. Optical Field of Regard Objects in the 180 hemisphere defined by the SoloHI external baffle plane (FOR) directly illuminate critical surfaces and requires special precautions within the
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instrument. Knowledge of the optical properties and size/distance of the protruding scattering surface are required to assess the impact. During Phase A, we modeled the impact of the solar array paddle and the RPW which intrude into this hemisphere. A complex SoloHI FOR was given to the spacecraft. Mass, Power, Envelope See Section 3.4.2
3.1.2 SoloHI Instrument Performance
The top-level instrument performance requirements for the SoloHI instrument are summarized in Table 3-2. The current SoloHI performance margin between the instrument capabilities at the beginning of life (BOL) and the related instrument requirements are in Table 3-3. This table includes the bandpass, telescope FOV, instrument spatial resolution, effective area, camera performance, and straylight rejection performance. The following sections describe the performance error budgets for the SoloHI effective area (3.1.2.1.1), instrument spatial resolution (3.1.2.1.2), and straylight rejection (3.1.2.1.3). The camera and instrumental straylight performance shown in Table 3-3. are required to meet the baseline science measurement photometric accuracy requirement.
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Table 3-2. SoloHI Instrument Performance Requirements and Capabilities
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Table 3-3. SoloHI Instrument Performance Requirement Margin
The requirement identification code in the following discussion is a unique identifier for every requirement in the SoloHI requirements documents, including the SoloHI Science Requirements Document (SRD), the SoloHI Instrument Requirements Document (IRD), and the SoloHI Instrument Specification (ISPC). This code is used to maintain requirement traceability, enable requirement tracking and track requirement verification during the SoloHI program. The first number corresponds to the level of requirement (3 for mission segment, 4 for instrument and 5 for subsystem level). The second number corresponds to the document where the particular requirement is described. The third number corresponds to the sequentially numbered requirement within the document.
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3.1.2.1 Telescope Scene Coverage
The square APS detector and the circular image projected on the detector plane are shown in Figure 3-3. This figure illustrates that the detector will capture continuous radial scene coverage of 40° and transverse scene coverage of 26.54°. Of course, the full transverse scene coverage of 40° is provided at the middle of the radial scene coverage from elongations of 11.7° to 33.3°. The pixels in the detector corners will be stripped from the images before sending the images to the SoloHI Payload Operations Center. The detector corners will be divided into blocks of 128 x 128 pixels. The pixel blocks outside the projected scene on the detector plane will not be downlinked.
Figure 3-3. Scene Image Projected on the SoloHI Detector 3.1.2.1.1 Effective Area
The effective area budget for the SoloHI telescope is presented in Table 3-4. The effective area at EOL sets requirements on the effective area at BOL and the throughput loss over the Baseline science mission due to radiation, contamination or material degradation. Similarly, the effective area at BOL sets the requirements on the telescope aperture area, detector quantum efficiency at BOL, and optics transmission loss at BOL. Currently, the margin on the effective area at BOL is 7.6%. Table 3-4. SoloHI Telescope Effective Area Budget Requirements and Capabilities
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3.1.2.1.2 Spatial Resolution
The SoloHI spatial resolution requirement and capabilities for the Full Frame and Inner FOV Subframe images is presented in Table 3-5. The image spatial resolution sets requirements on the on-orbit instrument spatial resolution with nominal pointing, the instrument pointing jitter and the instrument pointing stability. The error budget for the on-orbit instrument spatial resolution includes the baseline detector spatial resolution for the appropriate image binning, the baseline telescope spatial resolution before environmental tests, the ground spatial resolution shift before launch, and the post-launch spatial resolution shift from the ground resolution. The baseline telescope spatial resolution before environmental tests sets requirements on the optical spatial resolution for the following cases: Nominal Optics and Detector Alignment Nominal Optics and Worst-Case Detector Alignments Worst-Case Optics and Nominal Detector Alignments The margin on the optical spatial resolution with nominal optics and detector alignment for the baseline SoloHI optics design is 15.3% and 14.7% for the Full Frame and Inner FOV Subframe images respectively. The detector spatial resolution margin for both the Full Frame and Inner FOV subframe is near zero, because this resolution is dominated by the detector plate scale and is well known. The margin on the instrument pointing jitter and pointing windowed stability is 782% and ≥1900% respectively. The Solar Orbiter observatory pointing is much better than required for the SoloHI spatial resolution performance. The nominal optical spatial resolution margin for the baseline optics design was only 15.3% for the Full Frame image at the diagonal extreme in the rounded corner of the projected image on the detector. Table 3-6 presents the RMS spot radius margin for the baseline design across the image from the Boresight to the Edge center to the Corner edge for the Full Frame, Inner FOV Subframe and Radial Swath Subframe images. The optical spatial resolution margin will be greater than 50% and 25% along the radial centerline for the Full Frame images and Radial Swath Subframe images respectively. The detailed SoloHI image spatial resolution budget for the SoloHI Full Frame image, Radial Swath Subframe image, and for the Inner FOV Subframe images centered at 7, 15 and 20 Rsun are presented in Table 3-6. The Baseline telescope spatial resolution before environmental tests includes error budget components for the Nominal Design, Element Fabrication, Static Optics Alignment and Detector Alignment. The optical spatial resolution requirement with nominal optics and worst-case detector requirements in Table 3-5 is the combination (RSS) of the Nominal design and Detector Alignment requirements in Table 3-7. Similarly, the optical spatial resolution requirement with worst-case optics and nominal detector requirements in Table 3-5 is the combination (RSS) of the Nominal design and Detector Alignment requirements in Table 3-7.
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Table 3-5. SoloHI Spatial Resolution Budget Requirements and Capabilities for Full Frame and Inner FOV Subframe Images
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Table 3-6. Nominal Spatial Resolution Across the Image Field for SoloHI Full Frame, Inner FOV Subframe and Radial Swath Subframe Images
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Table 3-7. SoloHI Image Spatial Resolution Budget
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3.1.2.1.3 Straylight Rejection
The instrument straylight rejection requirement at the detector is defined to be a fraction of the combined brightness for the F and K coronae along a given line of sight captured in the image. Table 3-8 shows the variation of the straylight fraction. The stray light is 0.44x the coronal brightness at the detector inner FOV at the mission perihelion of 0.28 AU and 0.22x the coronal brightness at the detector outer FOV for observations at 0.70 AU. The instrument straylight rejection requirement at the detector is separated into the instrumental straylight rejection at the aperture and the lens barrel straylight rejection. There are multiple sources for the straylight that enters the SoloHI aperture. The following analysis includes solar disk light diffracted by the Solar Orbiter heat shield and the SoloHI forward baffles, and reflected straylight from neighboring spacecraft hardware that enters the A1 aperture after multiple reflections in the SoloHI interior. The straylight rejection at the A1 aperture by the SoloHI interior is divided into different off-pointing angle ranges from boresight that correspond to different lens barrel straylight rejection requirements. Table 3-9 defines the lens barrel straylight rejection requirement for each of the straylight incident angle ranges at the A1 aperture and describes the SIM interior surfaces that fall within each of these ranges. The SoloHI straylight from neighboring spacecraft structures includes reflected light from the Solar Orbiter solar arrays, reflected light from the closest RPW antenna, and the diffracted light from the SoloHI forward baffles that reflects off the light trap above the aperture and then reflects off of the SoloHI interior into the A1 aperture. All these neighboring spacecraft structures lie outside the SoloHI unobstructed FOV, but above the plane of the SoloHI structure top cover. Table 3-8. SoloHI Straylight Rejection Requirement as a Fraction of the Coronal Brightness
Table 3-9. Lens Barrel Straylight Rejection Requirement for Aperture Incident Angle Ranges
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The SoloHI straylight rejection requirements and capabilities for the Full Frame images are presented in Table 3-10. The smallest straylight rejection margin at any instrument distance from the Sun occurs for the straylight rejection of neighboring spacecraft structure at angles of 21° to 45° from the instrument boresight. This straylight margin varies from 90% at 0.28 AU to 181% at 0.70 AU. The minimum 11.1% margin for the SoloHI lens barrel straylight rejection is in the transition region of 20° to 21° from the instrument boresight. Although the margins for individual straylight error budget components can be smaller than 100%, the overall SoloHI straylight rejection estimates at the detector have margins greater than 380% for all of the cases in Table 3-10. The straylight rejection estimates for the Solar Orbiter solar arrays, the RPW antenna, and the reflected light of diffracted solar disk light by the SoloHI light trap are determined for the inner and outer FOV both at perihelion and 0.7 AU. Table 3-11 shows these individual straylight estimates for the nominal and worst-case SoloHI baffle alignments at the detector inner and outer FOVs at 0.28 and 0.70 AU. 3.1.2.1.4 Modeling Coronal Focal Intensities for Orbital Positions of 0.28 and 0.7 AU
An essential part of the photometric analysis is the determining brightness of the coronal scene. The SoloHI team modeled the brightness of the solar corona over the instrument field of view for various orbital positions. Figure 3-4 shows the results of this calculation. The F corona brightness estimate is an extrapolation from current LASCO and SECCHI coronagraph data and Helios data, while the K-corona brightness estimate is based on the Saito equatorial and polar model. The coronal brightness as a function of elongation was determined using a three- dimensional “ray trace” model to integrate the Thompson-scattered light along each line of sight. The shot noise equivalent coronal brightness from the scene is plotted for 1 and 10 second exposures. The shot noise curves for these exposure times are well below the brightness of the K-corona in the inner field; longer exposures are required in the outer field of view. Figure 3-5 gives the APS photoelectrons/pixel-second at the focal planes.
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Table 3-10. SoloHI Straylight Rejection Requirements and Capabilities for Full Frame Images
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Table 3-11. SoloHI Straylight Rejection Modeled Estimate at Detector Inner (Top Table) and Outer (Bottom Table) FOV
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Figure 3-4. F and K Equatorial Corona Brightness Estimate at Various Spacecraft-Sun Distances with Equivalent Photon Noise Shown for 1 and 10s Exposure Times
Figure 3-5. Count rates (electrons/s/pixel) at the instrument focal plane for the corresponding F and K Equatorial Corona. Curves follow the same legend given in the previous figure.
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3.1.2.1.5 SoloHI Photometric Analysis
The scientific objectives described in Section 2 require precise measurements of the K-corona signal. The photometric requirements for the various images and observing programs are given in the SRTM (Table 2-2).The K-corona measurement accuracy requirement for Full Frame and Radial Swath Subframe images is defined as a signal to noise ratio (SNR) for a binned pixel of ≥20 in the inner FOV and ≥5 in the outer FOV. This corresponds to a measurement error of ≤5% and ≤25% respectively. Additionally, the K-corona measurement accuracy requirement for Inner FOV Subframe images is ≥16, at 7, 15 and 20 Rsun within 0.29 AU, and ≥12 at 15 and 20 Rsun between 0.29 and 0.36 AU for a binned pixel. During Phase A, the team modeled the K-corona measurement (SNR) in the SoloHI field of view and calculated the minimum necessary integration times to achieve the SNR required by the science objectives. We then designed unique observing programs for each science objective and calculated the associated characteristics.
SoloHI signal is defined to be the equatorial K coronal brightness, Keq , integrated along the entire line of sight defined by the given pixel. The SoloHI signal is estimated by subtracting the brightness measurement for a single pixel, M, from the brightness background, B, that is estimated on the ground with post-processing software: S = M - B (3-1)
The absolute signal estimation error eS is given by: e = Sˆ - S = (Mˆ - M)-(Bˆ - B) S (3-2) = eM - eB where Sˆ and S are the signal estimate and true signal respectively, Mˆ and M are the measurement estimate and true measurement respectively, and Bˆ and B are the background model estimate and true background measurement respectively. eM is the absolute measurement error, while eB is the absolute background model estimate error. For purpose of this analysis, the standard deviation of both M and B were estimated and linearly summed. For the photometric analysis, we considered the dominant sources of error including the F- corona, the equatorial K-corona, the integrated star background, individual stars and planets, the instrumental straylight background, and the detector dark current. Sources of brightness measurement error are camera read noise, shot noise from both the scene and the instrumental straylight, and the dark current estimation error. The signal estimation error includes the background model estimate error. The brightness measurement captured over a given integration time is a function of the integration time, telescope effective area, and the signal flux rate. The signal flux that enters the instrument aperture is based on the pixel solid angle and on the instrument distance from the Sun. Summing images proportionately increases the observation SNR as does further binning of neighboring pixels. The dominant source of measurement error is the photon shot noise. Minimum integration time that will satisfy the SoloHI photometric accuracy requirements are given in Table 3-12 and Table 3-13. Table 3-12 shows these derived integration times for the Full Frame, Inner FOV Subframe and Radial Swath Subframe images for the Perihelion, Near Perihelion observing cases. Similar results are presented in Table 3-13 for the Far Perihelion,
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Southern and Northern observing periods. This calculation is based on the SoloHI camera photometric performance (read noise and dark current), SoloHI instrument straylight rejection, and the background model estimation error. In addition, the minimum integration time selects the single image exposure time such that the signal will remain within the linear full well for the camera. The assumed number of summed images and the pixel binning are indicated on each row of this table. The minimum integration time for the summed image must be less than the image cadence for the corresponding SoloHI baseline observing program. Similarly, the image cadence in the SoloHI baseline observing program must satisfy the image cadence requirement from the Baseline science measurement requirement in the SoloHI Science Requirements Document. The far right columns of Table 3-12 and Table 3-13 show the image cadence for the SoloHI Orbit I and Orbit II baseline observing programs and the SoloHI image cadence requirement. All of the above requirements are satisfied in this table with substantial margin. 3.1.2.1.6 Strawman SoloHI Observing Program Characteristics
The SoloHI team developed a strawman observing program to capture the required observations using the functionality in the baseline SoloHI instrument design. The required Full Frame, Inner FOV Subframe, and Radial Swath Subframe images are captured using the available SoloHI functionality and satisfy the photometric accuracy requirements for all pixels in the image. Table 3-14 presents the observing program characteristics for various programs and heights. The following discussion presents details relevant to the observing program and observing program characteristics table. The architecture and instrument functionality are flexible and readily accommodate a variety of observing programs. The camera ADC gain column of Table 3-14 describes the SoloHI pixel gain settings during the observing programs. The SoloHI APS allows on-chip gain selection of high gain or low gain mode independently for the top and bottom halves of the detector. The high gain mode reduces the read noise compared to the low gain mode, but also reduces the linear full well. The SoloHI observing program will use the high gain mode, if there is no row on the detector that will reach the linear full well during the single image exposure time. It is always used for the Inner FOV Subframe images and for the outer FOV detector half for all SoloHI Full Frame and Radial Swath Subframe images. The readout column of Table 3-14 describes the type of readout used to obtain the science data. The science program makes extensive use of a Progressive Scan (PS) readout mode. In this mode, the SoloHI camera executes a sequential reset and read of each detector row of the detector, such that the integration time for every row is the same. This mode is essentially identical to a rolling curtain shutter. For a particular image, the integration start and stop time for each row are slightly offset in time from the previous row.
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Table 3-12. SoloHI Minimum Integration Times for Perihelion and Near Perihelion Observing Cases
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Table 3-13. SoloHI Minimum Integration Times for Far Perihelion, North and South Observing Cases
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Table 3-14. SoloHI Baseline Observing Program Characteristics
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The SoloHI observing program also uses a Progressive Scan mode with Inner and Outer Loops (PS-IL and PS-OL respectively). This camera mode is similar to the Progressive Scan mode, but there are two loops defined over two continuous blocks of rows from n1 to ninner and from ninner + 1 to nouter. This camera readout will read out the inner loop block of rows for a defined number of loops, niL, read out the outer loop block of rows a single time, and then repeat. The integration time for every row in the inner loop is the same, and the integration time for every row in the outer loop is the same. However, the integration time for the outer loop is approximately niL times the integration time for the inner loop. The SoloHI camera baseline include independent readout of the detector top half and bottom half. The camera gain and camera mode can be independently selected for each half. A SoloHI Full Frame image or Radial Swath subframe include multiple 1Kx1K row blocks (usually ≤3) in the detector top half (inner FOV) and in the detector bottom half (outer FOV). The SoloHI Inner FOV Subframe images used for the Wave Turbulence observing programs only use a single row block. All of these row blocks must be rapidly read while maintaining the image cadence for the synoptic observing program. In addition, no pixels on any single row for each row block should reach the linear full well during the single image exposure time. In the inner FOV for instrument distances <0.33 AU, the integration time for each row block is <1.5 min. Since this is a short time in the Full frame synoptic cadence of 30 min or the Radial Swath image cadence of 5 to 6 minutes, each row block will be read out sequentially using the Progressive Scan camera mode. In the outer FOV for instrument distances of >0.42 AU, the sum of the summed image integration time for the individual row blocks will exceed the image cadence period. In this case, the row blocks in the outer FOV will be exposed simultaneously using the Progressive Scan mode with Inner/Outer Loops to fit within the image cadence period without saturating any pixels in the innermost row block which captures a larger measurement. If no row block would saturate during the single image exposure time, such as in the Full Frame outer FOV >0.5 AU, the row blocks in the outer FOV can be read out using the Progressive Scan camera mode with the same single image exposure time. Afterwards, a different number of single images can be summed for each row block in the SCE to meet the SNR. 3.1.3 SoloHI Design Heritage SoloHI fully exploits the scientific and technical heritage of the SECCHI/HI instrument on the STEREO S/C. As shown in Figure 3-6, the SECCHI/HI instrument has nearly identical design elements as the SoloHI. The door mechanism is very similar to that used on STEREO/SECCHI. As described below, the baffle design uses similar elements with modest modifications which accommodate the solar array scatter. The optical design of the SoloHI telescope is similar to the HI1 telescope on STEREO/SECCHI. Similar to SECCHI/HI, the forward baffle structure combined with the heat shield leading edge baffle and the diffracted light trap provide the necessary suppression of the incoming solar radiation. The verified performance of SECCHI/HI -15 baffle is ~10 B/Bs. The peripheral baffles prevent glints and diffuse scatter from the surrounding spacecraft surfaces lying within the SoloHI hemispherical field of regard. The APS detector selected for SoloHI is a better match to the Solar Orbiter radiation environment and stringent mass requirements than the E2V 42-40 CCD used on STEREO/SECCHI. The telescope field of view was chosen to be twice that of SECCHI/HI1. This enhanced field of view will image high latitude structures and more of the ecliptic plane as the mission progresses. The instrument will view ±20º to the north and south of the ecliptic plane beginning at 5º from Sun
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Figure 3-6. The SoloHI instrument Concept and Mechanical Layout is Similar to SECCHI/HI The MEL (see Appendix J) provides additional information on the SoloHI components and their associated TRL. SoloHI draws heritage from SOHO/LASCO and STEREO/SECCHI instruments and ongoing mission operations expertise. Team members were selected for their prior experience to perform this mission with the lowest risk approach. 3.1.4 Baffle Design, Performance and Analysis The baffle design specifically addresses the three dominant sources of scattered solar radiation reaching the instrument. These are: 1) direct incoming solar radiation, 2) diffracted solar radiation from the edge of the heat shield and 3) directly scattered solar radiation from the solar array and RPW antenna. The schematics in Figure 3-7 and Figure 3-8 show the essential components. The direct incoming solar radiation passes harmlessly over the top of the instrument. Smooth edges on baffles F1-F4 provide additional protection from the edge of the heat shield. Interior and side wall baffles are specifically designed to suppress the scattered radiation from the RPW antenna and the solar array. Peripheral baffles also provide additional protection for scatter from spacecraft components outside the instrument field of regard.
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Figure 3-7. Baffle system side view schematic. Dimensions show general scale but are not for reference. The shadow line shows the worst case edge of the sun at 2.05 degrees. F1 to F4 shade the lens aperture from diffracted radiation scattered at the edge of the heat shield. Interior and peripheral baffles protect the instrument from scattered radiation from the surrounding spacecraft
Figure 3-8. Baffle system top view schematic. Placement of interior baffles and front baffles shown. Dimensions show general scale but are not for reference. Details of side wall baffle and lens cavity baffle arrangement are not shown
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For direct incoming solar radiation, the design includes a 2.05° solar exclusion zone as shown in Figure 3-7. This zone is sufficiently large to encompass with margin the size of the solar disk at perihelion and the maximum offpoint. This direct solar radiation within the exclusion zone passes harmlessly over the top of the instrument. Diffracted radiation from the heat shield edge is captured in a diffraction light trap located within the A1 enclosure. The baffling is arranged so that the surfaces directly illuminated by the heat shield Figure 3-9. Calculated irradiance at the lens entrance diffracted light are not in the field of aperture. The calculation was done for the perihelion regard of the lens aperture. case, worst case off-point with a nominal heat shield Diffracted radiation from the heat position. Dashed lines show the lens entrance aperture shield is further attenuated by four additional edges. Both Glad and an NRL model were used to calculate the diffracted light at the entrance aperture for a number of positions. Figure 3-9 and Figure 3-10 show the relative irradiance as a function of distance along the A1 aperture plane for perihelion for the worst case pointing condition. Figure 3-9 shows the irradiance distribution for the nominal heat shield position. Figure 3-10 shows the irradiance distribution for the worst case heat Figure 3-10. Calculated irradiance at the lens entrance shield position. aperture. The calculation was done for the perihelion The SoloHI baffle is specifically case, worst case off-point with a worst case heat shield designed to reject scattered light position. Dashed lines show the lens entrance aperture from the directly solar illuminated solar array paddle and the RPW antenna. As shown in Figure 3-11, these assemblies significantly intrude on the hemispherical instrument field of regard. Although SECCHI/HI faced a similar situation with an instrument antenna on STEREO, the angular extent of the illuminated area within the instrument field of regard required two design modifications. First, we introduced a multistage baffle surrounding the lens. This baffle prevents the lens from viewing any surface directly illuminated by scatter from the solar array. Second, we changed the interior baffle surface finish from a Chemglaze painted finish to a darker Deep Space Black or Martin Black finish. A report from Sandia Laboratory describing the space qualification of deep space black is available on request. Finally, we optimized the geometry of the forward baffles, the interior baffles and the lens enclosure to suppress scattered radiation emanating from the solar array and
3-20 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 the RPW antenna. Zemax non-sequential ray trace was used to evaluate the resulting baffle design. Representative results are shown in Figure 3-12 and Figure 3-13. A summary of critical parameters used in the stray light analysis is given in Table 3-15. Requirements, performance and margins are presented in Section 3.1.2.1.3.
Figure 3-11. Solar array and RPW antenna geometry relatively to SoloHI and the heat shield. Figure generated from Zemax non-sequential ray trace analysis
Figure 3-12. RPW antenna illuminating the SoloHI instrument (only the brightest rays shown). Figure generated from Zemax non-sequential ray trace analysis
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Figure 3-13. Solar array illumination of the SoloHI instrument baffle (brightest rays only). Figure generated from Zemax non-sequential ray trace analysis Table 3-15. Stray Light Analysis Summary
Instrument Description Parameter Hand calculations for initial design. Design evaluated and modeled using Analysis performed Zemax non-sequential ray trace and FRED. 4% Lambertian, solar array oriented to avoid specular reflection directly Solar array scattering into SoloHI baffles. RPW Antenna 90% Lambertian. Baffle finish Measured BRDF for Epner Laser Black (~1% Lambertian). Instrument walls CFRP matte black finish, 4% Lambertian. 3.1.4.1 Modifications to the Proposal Design The design has been updated to accommodate the larger than expected solar array, the RPW antenna and the larger than expected uncertainty in the position of the edge of the solar heat shield. These updates included internal changes to the interior baffle geometry and changing the position of the F1 to F4 baffles. The length of the instrument was modestly increased to accommodate the changes the F1 to F4 edge design. 3.1.5 Structure Design and Analysis 3.1.5.1 Structure Characteristics The SoloHI instrument consists of a small rigid box structure comprised primarily of CFRP panels. The box structure incorporates the one shot door and provides an interface for the forward baffle ledge, the kinematic mounts, and the FPA enclosure as shown in Figure 3-14. The instrument linear baffle accommodation is achieved conveniently via a light-weight internal baffle box. The baffle box design and assembly features are detailed in the following section. The primary structure is M55J lay-up with cross orientation. Doublers of similar material will be used locally to stiffen the structure at the interface of the instrument mounts and door and
3-22 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 associated hardware attachment. The FPA assembly is mounted via a Ti-6AL-4V housing. The base of the box structure is mounted to semi-kinematic, flexural mounts similar to STEREO/SECCHI and other instruments. Rigid brackets have been designed to meet the height requirement for positioning the SIM off the S/C panel for alignment with the heat shield and to meet overall frequency requirement.
Figure 3-14. Instrument Exploded View 3.1.5.2 Structure and Instrument Assembly The CFRP structure walls are joined together using standard, qualified techniques for box enclosures. Panel clips and doublers will be used at the seams. The forward baffle ledge is considered a subassembly and attaches to the box structure to form an extension to the structure. The forward baffles are attached to the ledge via flexure clips and alignment provisions are incorporated for anticipated adjustments in the baffle tips during alignment testing. The internal baffle box is a self contained subassembly housing the interior baffles which have been treated with an optical dark black coating. Numerous handlings of the interior baffles can be minimized with this approach. The baffle box side walls have slots and bosses to accept attachment of the interior baffle tabs external to the box. Precision machining of the side wall bosses will provide self alignment of the baffle tips. Alignment provision in the integration of the system of baffles as a set will be accommodated with shims. The internal baffle box gets installed into the outer enclosure with few attachments and can be removed easily. The attachments for the baffle box are two bushings near the top rear for pinning and at the front along the bottom edge. The thin front plate of the baffle box is bent to act as a flexure.
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Figure 3-15. Instrument Internal Baffle Box The lens set mounts directly to a Ti-6Al-4V bulkhead at the back end of the structure on the front side. On the rear side, the separate FPA enclosure is attached via a series of fasteners. Access inside the FPA enclosure is readily available with removable panels. The kinematic mounts and door are attached with fittings to provide adequate load path for launch and on-orbit operations. 3.1.5.3 Modifications to the Proposal Design A primary change from the proposal is implementing the internal baffle box compared to attaching each baffle separately to the outer CFRP structure wall with flexure clips. Due to the number of interior baffles required to reduce stray light to acceptable levels, assembly of the baffles would have been excruciating to integrate and align individually in this manner. The CTE mismatch between the aluminum baffles and CFRP structure is now simplified by systemizing the approach. Fewer parts are required, and overall, a more cost effective and exciting solution prevails. The number of mounting feet was increased from three to four to meet the fundamental frequency requirement of 140 Hz. In addition, the kinematic mounting configuration was altered to achieve proper balance between frequency and thermal distortion. The original 3 blade, 2-2-2 degree of freedom configuration is now a fixed mount-2 blade-post configuration providing a 3- 2-2-1 degree of freedom arrangement. A mass increase of approximately 0.2 kg has been realized with this modification. 3.1.5.4 Structural Analysis A SoloHI finite element model has been created to evaluate fundamental frequency and support trade studies. Modeling techniques for the model are standard and material properties are consistent with as measured or are selected to provide additional margin. Table 3-16 shows the first six frequency modes with > 5% mass participation of SoloHI on the four kinematic mounts integrated with the two interface brackets. Figure 3-16 shows the fundamental mode of SoloHI in this configuration. Modeling of the internal baffle box attachment inside the structure reveals ideal stiffness with frequency of the system at 400 Hz and thermal stresses below 8 ksi for a 120°
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Figure 3-16. Structural Deformation Plot Showing Primary Lateral Z Frequency 3.1.6 Preliminary Thermal Analysis Preliminary thermal analysis for the SoloHI instrument module (SIM) was completed during Phase A. The SIM thermal model was developed and used to evaluate the FPA radiator sizing and the operational/survival heater average power to maintain the SIM component temperature limits. During Phase A, the SoloHI team has evaluated the following thermal cases: Hot Operational, Cold Operational, Cold Survival, and Hot Failsafe Survival based on steady-state temperatures at 3.5° from the nominal sun-pointing orientation. Further optimization of the thermal geometry will be required to reduce the heater power while maintaining the instrument temperatures. This optimization is expected to include heater locations, adjustment of thermal finishes, cold finger geometry, blanket arrangements and radiator size. The preliminary analysis shows that the overall design concept is sound and will not require extensive changes. The
3-25 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 analysis showing the survival in the off-pointing case was presented in the IDSR and has not been updated for the final CSR. 3.1.6.1 Instrument Thermal Model Description Figure 3-17 shows the SIM thermal model created in Thermal Desktop during Phase A, while Figure 3-18 describes the surface material or finish, while Table 3-17 lists the surface optical properties (absorptance and emittance) at BOL and EOL. SoloHI is shielded from direct sunlight by a spacecraft heat shield and is attached to and thermally isolated from the spacecraft deck by titanium feet. A composite door hinged on one side of the structure is opened 240° when the instrument is operational. The CFRP structure is covered with Black Kapton MLI blankets. The SoloHI Camera Electronics (SCE) Box is modeled as a five sided structure with black Kapton blankets on its four smaller sides and Z93 white paint on the large side to act as a radiator.
SCE Box FPA Enclosure Baffle Cover Interior Baffles Forward Baffles
FPA Radiator
Figure 3-17. SoloHI Thermal Model
Blue = Black Kapton MLI Purple = Black Anodized Aluminum Yellow = Z93 White Paint Gray = Epner Laser Black
Silver Teflon MLI
Figure 3-18. SoloHI Thermo-Optical Properties
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Table 3-17. SoloHi Optical Properties BOL EOL Material or Finish α ε α ε Black Kapton MLI 0.90 0.86 0.92 0.88 Black Anodized Aluminum 0.70 0.86 0.80 0.82 Z93 White Paint 0.17 0.94 0.22 0.88 Epner Laser Black 0.95 0.84 0.95 0.84 Silver Teflon MLI 0.11 0.75 0.24 0.74 The SoloHI team developed a detailed FPA model to describe the heat flow between the APS detector, the Detector Readout Board (DRB), the cold finger, and the FPA radiator. The FPA enclosure is modeled as a five sided titanium box covered with black Kapton MLI blankets. The enclosure is covered with black Kapton blankets and is connected to the back end of the baffle structure. The FPA enclosure houses the detector assembly that consists of the detector, Kovar chip carrier, a 6063 aluminum cold finger, DRB with titanium support frame, and a titanium cold finger mount. The detector is bonded directly to the Kovar carrier that is bolted to the cold finger. The cold finger extends through the titanium mount, the enclosure wall, and the radiator mount. The cold finger is bolted directly to the radiator. The Kovar carrier is electrically connected through Kovar pins to the Thermal Isolation Package (TIP). The TIP is connected using two flex circuits to the DRB that is supported by a titanium frame off the APS mount. The APS mount is bolted to the radiator mount through the enclosure wall. The detector, Kovar carrier, cold finger, APS mount, DRB, TIP, and support frame all radiate to the interior of the FPA enclosure. The radiator is modeled as a multi-node surface with a direct coupling to the cold finger. Black Kapton MLI blankets span the rear and bottom gaps between the radiator, FPA enclosure, and the instrument structure. The forward and top gaps between the radiator and the SIM structure are spanned by Silver Teflon blankets since these surfaces are directly illuminated by the solar disk in failure off-pointing cases. There is no direct solar loading on the radiator in the nominal operating cases. The radiator is coated with Z93 white paint on the front side and covered black Kapton MLI on the back side. The interior baffles are finished with Epner Laser Black paint. The forward baffles are black anodized aluminum on the surfaces facing the sun and Epner Laser Black paint on the surfaces facing the lens assembly. The baffle cover is also black anodized aluminum. 3.1.6.2 Environment Thermal Model Description The SoloHI environment is defined by a simplified spacecraft thermal model that assigns worst case boundary temperatures to the spacecraft surfaces. The top surface of the spacecraft is assumed to be a radiating surface for the electronics boxes mounted below and is coated with white paint. The solar arrays and the solar heat shield are also modeled using boundary nodes. The side of the spacecraft heat shield facing the instrument is covered with MLI with an outer coating of vapor deposited aluminum (VDA).
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3.1.6.3 SIM Thermal Analysis Summary Thermal analysis was conducted for three cases: Hot Operate, Cold Operate, and Cold Survival. The assumptions for the orbits, environments, and spacecraft boundary temperatures are found in Table 3-18. Power dissipations for the SCE box, the DRB, and the detector are found in Table 3-19. Table 3-18. SoloHI Thermal Case Assumptions
Solar Heat Shield Spacecraft Solar Array Solar Array Case Orbit Constant Temperature Temperature Temperature Rotation (W/m²) (°C) (°C) (°C) Hot Operate 0.28 au 28,200 70° 137 40 270 Cold Operate 0.92 au 1320 0° -6 0 30 Cold Survival 1.0 au 610 0° -22 -20 -20 Table 3-19. SoloHI Power Dissipations
Power Dissipations Hot Operate (W) Cold Operate (W) Cold Survival (W) APS Detector 0.16 0.16 0.00 Detector Readout Board (DRB) 0.18 0.18 0.00 SoloHI Camera Electronics (SCE) 5.42 5.68 0.00 Table 3-20. SoloHI Temperature Predictions
Hot Operate (°C) Cold Operate (°C) Cold Survival (°C) SIM Subsystems Predict Limit Margin Predict Limit Margin Predict Limit Margin Lens Barrel Assembly -20.0 30 50.0 -20.0 -30 10.0 -30.0 -50 20.0 SCE Baseplate 30.0 40 10.0 -9.2 -40 30.8 -39.0 -50 11.0 Detector Readout Board -44.3 -20 24.3 -44.7 -55 10.3 -54.3 -65 10.7 FPA Enclosure, Top -63.3 60 123.3 -73.4 -110 36.6 -89.7 -110 20.3 APS Detector -65.4 -55 10.4 -74.7 -85 10.3 -91.6 -110 18.4 FPA Radiator -68.1 100 168.1 -77.7 -130 52.3 -93.5 -130 36.5 SIM Structure -80.8 to 28.4 60 30.2 -104.6 to -65.4 -125 20.4 -105.2 to -75.6 -125 19.8 Forward Baffle Assembly -98.9 to -91.6 60 151.6 -108.2 to -104.0 -120 11.8 -108.9 to -104.1 -120 11.1 Interior Baffle Assembly -109.5 to-64.1 60 124.4 -109.5 to -77.1 -120 10.5 -110.0 to -86.0 -120 10.0 Light Trap Baffles -70.8 to -67.2 60 127.3 -88.0 to -84.0 -120 32.0 -98.1 to -95.7 -120 21.9 Table 3-20 shows the temperature predictions of various components for the Hot Operate, Cold Operate and Cold Survival cases. All of the SoloHI components are within their limits for all cases. Table 3-21 shows the power estimates of the heaters for the above cases. In the Hot Operate case, the SoloHI operational heater average power estimate is 1.53 W to maintain the temperature limits of the lens barrel assembly, the operational amplifiers on the DRB, and the interior baffles. The SoloHI operational heater average power estimate is 2.40 W for the Cold Operate case to maintain the temperature limits of the lens barrel assembly, APS detector, the operational amplifiers on the DRB, and interior baffles. In both cases, most of the heater power (1.30 W and 1.58 W for the Hot and Cold Operate case respectively) is used to maintain the
3-28 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 optical alignment of the lens barrel. A small amount of power is required to satisfy the internal baffle temperature limit that maintains the baffle alignment. The operational heater to control the APS detector temperature is only used during the Cold Operate case and is mounted to the cold finger underneath the detector Kovar tub package. The operational heater to control the operational amplifier temperature on the DRB is mounted to the DRB support structure. The SoloHI survival heater average power estimate is 5.90 W for the Cold Survival case. The same heaters are active for the Cold Survival case as for the Hot Operational case, with the addition of the SCE baseplate heater. Table 3-21. SoloHI Heater Power Estimates
Heater Power Estimates (W) Heater Locations Hot Operate Cold Operate Cold Survival Lens Barrel 1.30 1.58 1.40 Cold Finger 0.00 0.42 0.00 DRB Support Structure 0.13 0.30 0.56 Internal Baffle Assembly 0.10 0.10 0.27 SCE Baseplate 0.00 0.00 3.67 Total 1.53 2.40 5.90 3.1.7 Electronics 3.1.7.1 Electronics Description The SoloHI electrical block diagram in Figure 3-19 describes the electrical architecture for the SoloHI instrument and the external interfaces between SoloHI and the Solar Orbiter spacecraft. The SoloHI electronics includes the SoloHI Camera Electronics (SCE), the Detector Readout Board (DRB), the SoloHI Power Supply (SPS), the APS detector in the Focal Plane Assembly, the operational heaters, and telemetry points (for temperature, voltage, current, and door position measurements) for the SIM and SPS. The spacecraft directly provides switched power for SoloHI‟s survival heaters and for a paraffin actuator heater circuit to facilitate instrument door deployment. In addition, the spacecraft monitors a small set of temperature sensors near or on the SoloHI instrument. The 4Kx4K APS detector biasing, readout and data formatting is performed by the camera electronics, described in Figure 3-20, located within the SIM. The Detector Readout Board (DRB) is located near the APS Detector and provides the addressing, control, bias and reference voltages required by the detector. In addition, the DRB provides for the signal conditioning and conversion from analog to digital signals levels for the pixel signals. The SoloHI Camera Electronics (SCE) box is mounted to the aft face of the Focal Plane Assembly (FPA). The SCE provides timing to the DRB using pixel clocks, APS row and column addresses, and other timing signals required for the APS detector readout. Digital pixel information received from the DRB is converted to digital data and stored in local memory on the SCE. The SCE houses the Processor Card (PRC) and the Camera Electronics Card (CEC). The CEC communicates with the PRC for control and telemetry over a memory-mapped interface. Command functions have been identified to include readout, calibration and test modes. Telemetry will include digital telemetry associated with the detector readout process.
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Pixel information is transferred to the PRC using flight software during execution of the pixel- processing algorithm. Power for the camera readout electronics is derived from the Spacecraft operational power feed using isolated DC-DC converters located in the SPS. These voltages are expected to include a 3.3V digital supply and a Quiet ±5V analog supply. Additional power regulation will be included on the SCE to provide the APS detector and analog signal chain with the required bias, reference and operating voltages. The APS detector can be read out in two modes: a global reset mode and a progressive scan mode. In global reset mode, the FPGA on the CEC will control the global reset signal followed by selectively reading out a number of rows. In progressive scan mode, the FPGA resets individual rows at a time and, after a delta time, reads the individual row. The upper and lower halves of the stitched array detector readout can be operated independently. The combined pixel readout rate required is 4 Mpixel/s with an 8 Mpixel/s goal. The camera readout electronics conceptual design has investigated the use of a rigid-flex packaging approach to optimize the signal routing between the DRB and SCE. To reduce program risk, the camera electronics breadboard (CEBB) in Figure 3-21 was developed using the SoloHI analog camera electronics design and was used to evaluate the Sandbox VI detector in functional, performance and characterization tests performed in air and vacuum during the Extended Phase A.
Figure 3-19. System Electrical Block Diagram
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OFFSET & GAIN 1.5V SCE PIXEL A15..0 ADC UPPER WE CS PIXEL OE ADC LOWER D31..0 INTR INTF PC
25MHz UPPER ROW/COL, CLOCKS Interface DRB Processor Interface Processor 50MHz
OSC LOWER ROW/COL, CLOCKS CLOCK 50MHz CAMERA INTF DIST 25MHz 25MHz RTAX 2000 A14..0 CAMERA Local 25MHz RAS/CAS/WE INTF SDRAM CS6..1 ADC 64Mbx48 INTF DQM6..1 ±5V SCE BIAS V 256MByte DQ47..0 REF V PRS EDAC
CLK/CKE
MEMORY CNTL MEMORY SIM Power & CVTR SIM
DOUT Door Encoder TLM TEMP11..1 Analog MUX DIN 12 Bit ADC IO V&I13..1 CS3..0, SCLK Power Control 7..0
Signal Cond. RS422 ADC INTF ADC
Power 3.3V SCE Condition ±5V SCE
Figure 3-20. Camera Electronics Card (CEC) Architecture
Figure 3-21. SoloHI Camera Electronics Breadboard (CEBB) The PRC design shown in Figure 3-22 has baselined the LEON3FT-RTAX FPGA as the processor. The PRC will capture the images that are readout by the CEC in the 256 Mbyte SDRAM memory with Parallel Reed Solomon error correction. Flight software processes and compresses the image data, and formats the resulting science data for eventual transmission to
3-31 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 the spacecraft over one of the two SpaceWire interfaces. Instrument telemetry is also formatted for transmission to the spacecraft using the SpaceWire interface. Commands and timing information for the SoloHI instrument are received over the SpaceWire from the spacecraft and are parsed by the flight software for appropriate actions. The processor card within the SCE also supports boot PROM and NV memory for application program storage, loading and updating. In addition, analog to digital signal conversion capability is included for voltage, current and temperature telemetry points originating from the SIM and from within the SCE and SPS.
TxDS_PRI TxDS RxDS_PRI RxDS INTR TxDS_SEC LVDS RxDS_SEC EN CameraInterface
Spacewire A15..0 BA15..0 SPACEWIRE WE BWE CS BCS A14..0 OE BOE RAS/CAS/WE D31..0 BUFFERS BD31..0 SDRAM CS6..1 64Mbx48 DQM6..1 256MByte 1.5V DQ47..0 CONTROLLER MEMORY PRS EDAC RESET 25MHz LEON3FT- RTAX 50MHz 2000 A16..0 PRS 25MHz WE4..0 MRAM CID7 25MHz CLOCK 25MHz 2Mx8 (2) CS7..0 DIST
4MByte OE CONTROLLER MEMORY 32KB Boot D39..0 Prom RBSY7..0 RESET
RESET 25MHz 1.5V 1.5V 3.3V SCE REG Serial Debug RS232
NonFlt UART
Figure 3-22. Processor Card (PRC) Architecture Margins have been calculated for the SDRAM, boot and NV memory storage area and FPGA usage and devices selected with 2X margin to facilitate size, weight and power estimates. Custom sections of Hardware Description Language (HDL) code for the LEON3FT-RTAX CID7 device have been identified and a development plan established to accommodate the image pixel interface, SPI port analog to digital converters, start up boot ROM, default GPIO, clock distribution and SDRAM error correction. The SPS design in Figure 3-23 contains two power supply cards, the Power Control Card (PCC) and the Power Converter Card (PCNV), that accepts primary and redundant spacecraft operational power and power control signals, and provides power control select status back to the Spacecraft in addition to the secondary power feeds required by the SCE and the SIM. The SCE requires a 3.3V digital and a quiet ±5V analog isolated supply. The PCNV provides isolated power switching for operational heaters on the SIM. The flight software on the LEON3FT FPGA controls the power switching.
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Spare SSR
Spare SSR
Lens Barrel Htr SSR
Baffle Htr SSR
Door Hinge Htr SSR
Door Mech Htr SSR TBD Door LED Door LED SSR CM FILTER HYBRID SOURCE SIM
28V PWR
- TLM TLM CMD RELAY SC SSR Control7:0 - TLM
PRI INRUSH EMI TLM LIMIT FILTER
PWR 28V - 3.3V DIG CM 3.3V CMD SC FILTER
- RELAY HYBRID TLM DAMP
FILTER SC OPERATIONAL OPERATIONAL SC POWER RED ±5V ISO CM ±5V
FILTER HYBRID SCE ELECTRONICS SCE
Figure 3-23. SoloHI Power Supply Cards The SPS will house two card in a vertical orientation in its enclosure, shown in Figure 3-24. The SPS will be mounted on the +Yopt spacecraft deck exterior under the SIM. The SPS enclosure wall thickness is currently 3.5 mm to mitigate radiation total ionizing dose (TID) to ≤ 100 krad (Si) with a 3X margin goal. Internal interconnect between the processor and power supply card is accomplished with a flight quality motherboard/daughter board style connector.
Figure 3-24. SPS Enclosure
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3.1.7.2 Modifications to the Proposal Design The proposal design for the Processor Control Card (PCC) located within the SCE baselined the use of the AT7913E SpaceWire Remote Terminal Controller (RTC) ASIC developed by ESA and fabricated by Atmel. During the Phase A effort, a trade study (see Table 3-22) was conducted assessing the complexity differences between the AT7913E RTC device versus the Actel RTAX FPGA based LEON3FT-RTAX RTC device. The proposal design for the PCC was estimated to include the AT7913E ASIC in addition to two other FPGA devices to accomplish memory control and image processing. The main basis for the amount of complexity external to the AT7913E in the proposal design was driven by the assumption to compress the image data with a hardware based algorithm and the need to store the images in local memory that is larger than the memory addressing capability of the AT7913E. The CIE electronics design was modified to accommodate the interfaces of the Sarnoff APS device. Table 3-22. Phase A RTC Trade Study
Area of Interest AT7913E SpWRTC ASIC LEON3FT-RTAX FPGA Performance, MIPS 34 MIPS, CPU@50MHz, Test June 20 MIPS, CPU@25MHz 2008 Performance, MFLOPS 2.5 MFLOPS 4 MFLOPS Spacewire, Tx & Rx 200 Mbps 50 Mbps High Speed FIFO 295 Mbps Not Supported, Standard Interface Configuration Memory Controller 32 MB SRAM Only, No Wait States 1 GB SDRAM, SRAM with Wait States TID 300 krad (Si) 50 krad (Si) for Timing 300 krad (Si) Functional (1%) SEU >25 MeV/mg/cm2 Logic >37 MeV/mg/cm2 Logic >3 MeV/mg/cm2 Memory Cell 2.3 e-9cm2/bit Memory Cell SEL >80 MeV/mg/cm2 >104 MeV/mg/cm2 Power 150 mW 500 mW Voltages 1.8, 3.3V 1.5, 3.3V Silicon Quality Level ATMEL Space Level B, Assume QML- QML-V Eq. Flow, In Process of SMD Q Eq. Flow Approval Package MCGA349 CQ352 or CG624 The LEON3FT-RTAX design concept considered during the trade study provides for the direct control of SDRAM memory in amounts required by the instrument. Adding a custom interface to the LEON3FT-RTAX core permits image pixel information to be stored into the SDRAM using the AMBA DMA resources within the LEON3FT IP Core. In addition, Phase A investigations determined that a LEON3FT-RTAX Sparc V8 with a Floating Point Unit would have the processing capability with margin to perform the image compression using software based algorithms. The resulting reduced complexity of the LEON3FT-RTAX RTC design permits a reduction in PWB size and power for the PCC enabling closer compliance with spacecraft power and mass accommodations. The SpaceWire links core used in the LEON3FT-RTAX are similar to those used in the AT7913E, however the maximum link data rates are somewhat lower in the LEON3FT-RTAX due to the FPGA implementation. The trade study investigated the use of a 10
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Mbps command link rate and a 50 Mbps telemetry link rate and determined these would satisfy the instrument data bandwidth requirements if suitable for the spacecraft SpaceWire network. 3.1.7.3 Heritage EEE parts selection for the conceptual designs of the instrument camera readout and control electronics are based on qualified parts listed on the DSCC QPL that meet the 2X radiation requirement and often the 3X goal. The ADC currently being considered for the detector readout board is available as a 300 krad (Si) class V part manufactured by STMicro. Operational amplifiers have been identified and baselined for the conceptual design that have 100 krad (Si) TID tolerance and are available as class V devices. The LEON3FT-RTAX RTC has flight heritage on the Indian Chandrayaan-1 mission and processor peripheral functions and memories are either 100 krad (Si) class V or Q devices and in some cases are mitigated for SEU effects such as for SDRAM devices. The packaging concept for the SCE has been used by the NRL on past programs including HREP, TACSAT4 and Upper Stage. 3.1.8 Mechanisms 3.1.8.1 Mechanisms Description The instrument incorporates one mechanism, which has heritage from the SECCHI/HI. The SoloHI “one-shot” door is actuated by applying power for a predetermined time, and uses a Starsys paraffin actuator with microswitches to sense open/closed positions. The door lid is a rigid lightweight CFRP panel. The door is restrained by hinges on one edge and clamped by a latch mechanism on the other edge. The latch is held in place by a locking arm. The actuator releases the locking arm and the spring-loaded hinge then opens the door. The actuator has margin to release the locking arm, and the spring has margin to overcome friction and force the door fully open. No damping was required on the SECCHI/HI door as the mass of the door is small. Low outgassing lubricants are used. 3.1.8.2 Modifications to the Proposal Design No design changes from the proposal were made. 3.1.8.3 Heritage The design for the SoloHI door mechanism follows the design concept of the SECCHI/HI door. 3.1.9 Optical Design 3.1.9.1 Optical Description The SoloHI telescope optical layout is shown in Figure 3-25. The preliminary lens prescription is given in Table 3-23. The aperture stop is located in front of the lenses, in order to minimize the illumination of the lens by external surfaces. The incident radiation passes through the lens and is focused onto the APS detector. The telescope FOV is shown in Figure 3-26 and the geometric spot sizes are shown in Figure 3-27. The lens image quality is well optimized over the 24 circular field of view. The detector corners are largely unobscured. The geometric contribution dominates the instrumental point spread function. A preliminary spatial resolution performance budget is presented inTable 3-5. Similar to the SECCHI HI1 design, the bandpass is determined by a combination of long and short wavelength cutoff thin film filters placed on one of the lens elements. A critical parameter is the suppression of stray light from bright objects outside the field of view. The off-axis rejection was modeled by both NRL and CSL. The results are shown
3-35 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 in Figure 3-28. The first element of the design uses radiation tolerant glass. The lenses will be mounted in a barrel using an approach similar to STEREO/SECCHI. A preliminary look at the tolerance build-up using the Jen-Optik precision tolerances showed almost no change in overall image quality. 17:20:23
910 12 15 8 11 67 4 5 1 3
10.42 MM HI SO pup16x10 recul_det Scale: 2.40 CSL 25-Mar-10 Figure 3-25. SoloHI Lens Layout
Table 3-23. SoloHI Lens Prescription
Label RDY THI Glass Stop L1 37.165244 3.235618 LAK9G15_SCHOTT 879.288411 6.978516 L2 -32.801943 4.5 SF4_SCHOTT 87.701668 3.919768 L3 -66.971876 5 NLAF33_SCHOTT -33.071909 1 L4 49.301646 8 NSK16_SCHOTT -59.075697 23.214698 L5 -24.987039 3.100180 F5_SCHOTT -95.479699 15
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Figure 3-26. SoloHI telescope field of view
0.4900 0.6000
OBJ: 0.00, 0.00 (deg) OBJ: 0.00, 5.00 (deg) OBJ: 0.00, 10.00 (deg) 0.7400 100.00
IMA: 0.000, 0.000 mm IMA: 0.000, 4.935 mm IMA: 0.000, 9.940 mm
OBJ: 0.00, 15.00 (deg) OBJ: 0.00, 20.00 (deg) OBJ: 0.00, 24.00 (deg)
IMA: 0.000, 15.093 mm IMA: 0.000, 20.479 mm IMA: 0.000, 25.028 mm
Surface: IMA Spot Diagram
12/14/2011 Units are µm. Field : 1 2 3 4 5 6 RMS radius : 8.903 8.982 8.877 8.171 7.288 11.553 GEO radius : 14.784 18.347 19.504 19.416 17.746 38.844 SoloHI_CSL_23Nov2011_01.ZMX Scale bar : 100 Reference : Chief Ray Configuration 1 of 1 Figure 3-27. Lens Spot Diagrams as a Function of Field Angle
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1.E+00 0mm
1.E-01 1mm
1.E-02 2mm
1.E-03
1.E-04
1.E-05
1.E-06
1.E-07
1.E-08 Flux on detector [%] - lens barrel retainer diameter retainer barrel Fluxlens - impact detector on [%]
1.E-09 0 10 20 30 40 50 60 70 80 Incident angle [arcdeg]
Figure 3-28. Off-Axis lens stray light performance. The three curves show the sensitivity to the lens barrel geometry. The orange curve is the SoloHI baseline design 3.1.9.2 Modifications since the 2010 CSR The lens design was optimized for overall imaging performance over the larger 24 degree field of view. A study of the image quality variation with the lens fabrication tolerances was conducted. The Jen-Optik precision tolerances showed minimal degradation for a lens assembled with these tolerances. A draft of the lens specification is presently under revision in preparation for upcoming Phase B procurement activities. 3.1.9.3 Heritage The lens optical design is similar to designs developed for SECCHI/HI. The potted lens mounts are expected to be very similar to lenses provided for SECCHI/COR 2. 3.1.10 Detector System 3.1.10.1 Detector Description The APS detector is a 4096x4096, 10 μm device under development by SRI Sarnoff Corporation. The SoloHI team selected the 6T Pinned Photodiode (PPD) pixel format, shown in Figure 3-29. Selected APS performance requirements are given in Table 3-24. The pixel block diagram in Figure 3-30 shows the architecture of the device. Further details are given in the Sarnoff Phase A design report, the technology development plan and the contract statement of work (available by request). The 14-bit ADC is off chip along with other support functions (voltage regulation, LVDS drivers, bias generator, and readout generator). The individual pixels in the APS can be read out directly avoiding CTE related issues present in CCDs. The direct readout enables a shutterless operation, without the smearing that is present on SOHO/EIT and SECCHI/HI. The device will be operated in a combination of progressive scan and reset/read row operations to obtain the SoloHI images. Although the APS devices are not expected to exhibit flat-band shifts or any measurable CTE degradation, a modest dark current increase from radiation damage is expected. The Sarnoff minimal detector family has the following highly desirable characteristics:
3-38 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 proven radiation performance (>1 Mrad), well understood pixel design/trade-offs, and device performance extensively tested/documented in the open literature. The expected device performance meets the SoloHI performance requirements.
Figure 3-29. 6T PPD Pixel Design Schematic
Table 3-24. Selected APS Characteristics
Characteristic Nominal Value Format 4K x 4K, 6T pixel Pitch 10 μm Operating Temperature <-50º C Technology Jazz 0.18 μm, stitched Power <500 mW at 3.3 V ≥28% 500-700nm (Goal) QE ≥25% 500-700nm (Reqt) Radiation Tolerance >36 krad <60 electrons at low gain Read Noise <6 electrons at high gain Operation Mode progressive scan, row reset and read >86.4 K electrons at low gain Linear Full Well >19.2 K electrons at high gain >4 Mpixels/sec/port (Goal) Readout rate ≥2 Mpixels/sec/port (Reqt) Digital Controls Off-chip ADC Off-chip
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Figure 3-30. SoloHI APS block diagram. The individual pixel controls are addressable on a row by row basis During the phase A study period, we initiated a design/fabrication program with Sarnoff Research Corporation. The Sarnoff design team adapted the minimal pixel design to the SoloHI requirements. A switched MIM capacitor was added to the current Sarnoff pixel design (Figure 3-29). The resulting switched gain capability is an excellent match to the SoloHI requirements. The pixel design was rescaled to meet the SoloHI 10 micron requirement. SoloHI takes advantage of the minimal device readout frame on-chip double sample and hold circuitry. The resulting buffered signal simplifies the readout circuitry in the CIE and DRB. Representative devices with this pixel and the latest version of the frame design will be fabricated in the Sandbox VI foundry run presently underway at the Jazz foundry. The final device will be fabricated using the designs developed and tested in Sandbox VI. To reach the required array size with an active area of 4 cm x4 cm, a stitched device is required. Devices larger than this must be created by stitching patterns from a single reticule. Stitching is used routinely by semiconductor foundries to create large CMOS devices. Design rules and process steps are well documented and known. The general stitching process from a single reticule mask set is shown in Figure 3-31. During the Phase A study, Sarnoff developed a design concept for the SoloHI stitched array. The preliminary device layout is given in Figure 3-32 with patterns shown in Figure 3-33. As of this writing, a statement of work for the design, fabrication and test of the flight device is presently undergoing a competitive bid process. Preliminary pad layouts and a design of the flight packaging are given in the Sarnoff design report. The design report also includes further detail on the screening and testing of the flight devices.
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Figure 3-31. Stitched reticule geometry. This figure shows how stitching is used to construct a large scale array
Figure 3-32. SoloHI APS detector map showing the stitching blocks
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Figure 3-33. Stitching patterns. Preliminary results from the stitching study show that 5 stitching patterns (P1 to P5) are required to stitch the SoloHI array In September 2011, NRL awarded a development contract to SRI Sarnoff to fabricate the flight device. The first phase of the contract was to bring the device design to a preliminary design review level. The preliminary design review is scheduled for January 24, 2011. Sarnoff has closed a number of design issues during this design phase, specifically including the following items: • Started to define the function of 1k block enables • Worked on pad strapping and corner stitched pattern • Continued to develop Mk x Nk PCB engineering package and low insertion force socket • Continued to develop wafer probe card • Updated baseline Mk x Nk pixel design. • Continued to develop mix and match 1k x 1k and 4k x 4k stitched designs. • Added ESD and output short circuit protection. • Isolated top and bottom circuitry. Center stitch pattern will extend through pixel frame and pixels. • Added exposure control to the Mk x Nk imager. Control can be disabled if not desired. 1024 different exposures are possible for base progressive scan. A preliminary wafer layout for the flight device based on this updated design is shown in Figure 3-34.
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Figure 3-34. Preliminary 4kx4k device wafer layout with 1kx1k test imagers
3.1.10.2 Testing of the SoloHI Minimal Prototype SRI Sarnoff designed and fabricated a prototype SoloHI detector during the SoloHI Phase A study period. The prototype SoloHI detector was adapted from the Sarnoff Minimal pixel design by the SRI Sarnoff design team to satisfy the SoloHI requirements. The pixel design was rescaled to meet the SoloHI 10 µm requirement. The SoloHI detector uses the Minimal device readout frame. The resulting buffered signal simplifies the readout circuitry in the DRB and SCE. Representative devices with this pixel were fabricated in the Sandbox VI foundry run. The Sandbox VI device has been extensively tested by the NRL Solar Physics branch, the Goddard Detector Branch, and SRI Sarnoff. The test effort at NRL was lead by Greg Clifford, Dennis Wang and Clarence Korendyke. The test effort at Goddard was lead by Cheryl Marshall and Augustyn Waczynski. The testing at SRI Sarnoff was conducted by Jim Janesick. The device has proven to be relatively easy to drive with a range of voltages resulting in excellent performance that meets or exceeds the SoloHI performance requirements, described in Table 3.X. In some cases, the results have not been fully validated and may be subject to revision. Final results will be presented at the technology readiness promotion review scheduled for the February/March time frame of next year with radiation testing scheduled for mid-January at UC Davis.
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Table 3-25. Summary of Current Sandbox VI Performance Test Results
Characteristics Requirement Measured parameter
Full well High gain: >19.2k electrons ~35000 with high gain; Low gain: >86k electrons ~150000 with low gain Linearity at low gain looks very good.
Quantum efficiency ~25% average from 490 to740 Average wavelength response of the device is as expected.
Imaging Within 80% of the Nyquist limit, Extremely good. minimal blooming
Read Noise High gain: ≤5 electrons (at BOL) Low gain: ~1.8 electrons (GSFC) Low gain: ≤60 electrons (at BOL) High gain: ~25 electronics (NRL)
Dark current <0.476 electrons/pixel-sec at Satisfies the requirement at operating temperature at BOL, temperatures <-40 deg C. factor of 12 increase allowed for EOL.
Operation mode Progressive scan Measurements taken with progressive scan
Figure 3-35. SoloHI minimal device performance as measured by Jim Janesick. This shows a full well of 35k and 154k electrons with noise of ~4.8 and ~25 electrons respectively
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SoloHI minimal device photometric performance as measured by Jim Janesick is given in Figure 3-35. Photographs of the NRL detector test set are given in Figure 3-36. A representative photon transfer curve for a single pixel taken at NRL is given in Figure 3-37. Representative full well and read noise for a row of pixels taken at NRL are given in Figure 3-38 and Figure 3-39. The Goddard Detector branch obtained the measurements presented in Figure 3-40 to Figure 3-44. Dark current measurements are given in Figure 3-40. The dark current performance for devices operated at 3.6V meets the requirement at temperatures <- 40 degrees C. Figure 3-41 shows dark current images taken at various temperatures for two devices under test. Relative wavelength response of the SoloHI devices is given in Figure 3-42. Preliminary linearity plots for high and low level signals are shown in Figure 3-43 and Figure 3-44. Figure 3-36. NRL CEBB custom camera built by Silver Engineering (top). NRL test chamber with optical port being The preparations for the proton used for minimal device testing (bottom) radiation testing on the Sandbox VI devices are well underway. The proton radiation test plan has been written. The test will be performed at the UC Davis by the Goddard Detector branch with 63MeV protons. Two devices will be exposed. One of the devices will be exposed at the flight dose and the other at 2x the flight dose. The devices during the test will be cold, biased and operated in a flight like manner. Performance characterization will be conducted at discrete intervals during the exposures. A sector analysis was performed to estimate the radiation level at the detector using the shielding due to the focal plane assembly and instrument materials and geometry. The sector analysis resulted in a TID estimate of 17.8krad using the SoloHI geometry. The effective NIEL fluence is 4.76e10 for 10 MeV protons and 1.11e11 for 63MeV protons. The NASA guidelines require a factor of 2 (100% margin) to account for uncertainities in determining the radiation environment.
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These test levels (36 krad for TID, 9.52e10 for 10 MeV proton NIEL fluence, and 2.22e11 for 63 MeV NIEL fluence) are well within the expected tolerance of the device.
Figure 3-37. Photon transfer curve for the Single Pixel readout of the Sandbox VI Detector
Figure 3-38. Read Noise for a Row of Pixels for the Sandbox VI Detector at High Gain using the Camera Electronics Breadboard (CEBB)
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Figure 3-39. Linear Full Well for a Row of Pixels for the Sandbox VI Detector at High Gain using the Camera Electronics Breadboard (CEBB) SoloHI Dark Rate 100.000
Detector #2 (4.5V) Detector #3 (4.5V) 10.000 Detector #4 (3.6V) Detector #5 (3.6V)
1.000
0.100 Dark Rate (counts/second)
0.010
0.001 190 200 210 220 230 240 250 260 270 280 290 Temperature (K)
Figure 3-40. Dark Current Rate for Various Sandbox VI Devices over the Temperature Range of -73 to +7 deg C and over the Voltage Range of 3.6 to 4.5 V
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Figure 3-41. Dark Images for Various Sandbox VI Devices over the Temperature Range of -73 to -33 deg C and over the Voltage Range of 3.6 to 4.5 V SoloHI Relative Response 300
250
200
150
Detector #3 (4.5V) Detector #2 (4.5V)
Relative Response 100 Detector #5 (3.6V)
50
0 400 500 600 700 800 900 1000 1100 Wavelength (nm)
Figure 3-42. Relative Wavelength Response for Various Sandbox VI Devices over the Bandpass of 400 to 1100 nm and over the Voltage Range of 3.6 to 4.5 V, measured by GSFC Detector Branch. Quantum Efficiency is roughly 25% over the 490 to 740 nm bandpass. Final calibration has not been applied yet
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Figure 3-43. Linear Response to a High Signal for Sandbox VI Devices at High Gain, measured by GSFC Detector Branch
Figure 3-44. Linear Response to a Low Signal for Sandbox VI Devices at High Gain, measured by GSFC Detector Branch
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3.1.10.3 Modifications to the December 2010 CSR Design The device design has been further refined since December 2010 but no significant modifications have been made. 3.1.10.4 Heritage The device is presently considered to be a TRL 5+ device with extensive prototyping carried out in the Sarnoff Sandbox development program. Various devices developed under the Sandbox program span the range of SoloHI device requirements. A pixel design meeting the SoloHI requirements was developed for the SoloHI minimal array on the Sandbox VI foundry run. Devices from this foundry run will demonstrate at a prototype level the required device performance (TRL-6) prior to the SoloHI PDR. Projected testing of these devices includes: imaging and photometric characterization and radiation tests. Acceptance testing will be conducted under vacuum and at the anticipated operating temperature. 3.1.11 Focal Plane Assembly (FPA) 3.1.11.1 FPA Description The FPA provides the SoloHI APS with physical mounting, optical positioning, electrical connection and thermal cooling. The SoloHI APS arrangement is shown in Figure 3-45. The FPA design is an updated design from the SECCHI/COR2 FPA. Design modifications were introduced to specifically address the larger detector, the APS control electronics, a warmer operating temperature and a shorter distance to the radiator plate. As shown in the exploded view in Figure 3-46, the detector incorporates a minimal number of parts and thereby reduces design/assembly cost and technical risk. The Kovar-mounted APS will be attached to a cold finger. Cooling will be accomplished using a radiator directly connected to the cold finger and passively radiating the heat to deep space. A 10C temperature drop between the radiator and the detector is expected based on the SECCHI performance. No difficulties obtaining temperature <-45C are expected since its design features are based on the SECCHI/COR2 FPA that maintained the detector temperature at operational temperatures of <-70ºC).
Figure 3-45. FPA design concept. The figure shows the side mounting of the FPA in the SoloHI structure. Preliminary sizing of the Kovar mount and detector readout board are shown
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Figure 3-46. SoloHI focal plane assembly exploded view. The SoloHI focal plane assembly uses a minimal number of parts. The 3 boss detector mounting surface is not shown 3.1.11.2 Modifications to the Proposal Design The SoloHI team updated the proposal design. The mechanical linkage to the radiator was optimized. The detector readout board and associated support was added. Preliminary thermal and mechanical analysis showed that the design is robust and readily accommodates the detector and readout electronics thermal requirements. 3.1.11.3 Heritage The design is similar to the SECCHI/COR2 and LASCO FPA designs. 3.2 Software Definition and Management 3.2.1 Flight Software 3.2.1.1 Flight Software Architecture
The SoloHI flight software is hosted on the LEON3-RTAX FPGA and runs on the Real- time Executive for Multiprocessor Systems (RTEMS) operating system. The SoloHI flight software design is modular. The SoloHI flight software architecture is presented in Figure 3-47. The
SoloHI flight software Figure 3-47. SoloHI Flight Software Architecture modules are organized
3-51 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 under 4 software layers: Real-time Operating System: Primary interface to hardware (processor, memory controller) and external interfaces (Spacewire, Cameralink, and serial UARTs). Supports basic software functionality, including task creation/control/communications, interrupt control/handling, memory allocation/deallocation, and watchdog timer services. Foundation Class Layer: Provides definition of CCSDS packet, containers, and parameter tables. Handles time management with hardware clocks on processor card. Provides decompression of observing program, parameter tables and software updates. Application Services Layer: Provides task management/prioritization, task message routing with a software bus, and instrument time management/synchronization. Application Layer: Generic applications shall perform CCSDS packet creation and transfer to spacecraft SSR, software module/file validation, memory dumps, memory-to-memory transfers, and memory scrub using Error Detection and Correction (EDAC) hardware support. Mission-specific applications shall perform command ingest, housekeeping, instrument health, instrument control, image processing, and image scheduling.
The SoloHI flight software modules for Table 3-26. SoloHI Flight Software Modules each of the four SoloHI software layers are Acronym Flight Software Module listed in Table 3-26. The SoloHI flight software in the Foundation Class layer, OS Real-time Operating System Application Services layer, and Generic Mission-Specific Application Layer Applications was developed under the CI Command Ingest SECCHI instrument suite program and has HKP Housekeeping been used successfully over the last four IC Instrument Control years since the STEREO launch in October IH Instrument Health 2006. This code will have a high level of IP Image Processing reuse (90% reuse). This code will primarily IS Image Scheduling only have to be updated based on the Generic Application Layer change in the real-time operating system FE File Export from VxWorks 5.4 for SECCHI to the current RTEMS for SoloHI. The SoloHI MS Memory Scrub Mission-specific Applications will only SM Software Manager have a 50% reuse, due to the specific Application Services Layer software requirements for the SoloHI BT Background Task mission. Table 3-27 summarizes the SB Software Bus/Library functionality of the key SoloHI flight SCH Scheduler software modules. TM Task Manager Foundation Classes Layer SO System Objects UT Utilities
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Table 3-27. SoloHI Flight Software Tasks SoloHI Flight Functional Description Comment Software Tasks Task Manager (TM) Manage software tasks Start and stop tasks Command Ingest (CI) Accept commands from S/C Timekeeping, S/C status, instrument commands Software Manager Manages non-volatile Storage of parameters and code in non- (SM) memory volatile memory File Export (XF) Send science data to S/C Buffers Science data to S/C Instrument Health (IH) Autonomy rules Monitors instrument health and safety Instrument Control (IC) Camera controller Takes images Memory Scrub (MS) Memory scrub for SDRAM Memory scrub for SDRAM Housekeeping (HK) Send housekeeping packets Send housekeeping packets Image Scheduling (IS) Manages observing Tells Instrument Control (IC) when and schedule what images to take Image Processing (IP) Image summing, binning Lossless (Rice) and Lossy (H-compress and compression wavelet) image compression Figure 3-48 shows the hardware interaction between the SoloHI software modules, the electronics, and the SoloHI Software Bus module. The SoloHI design uses a proven real-time operating system (RTEMS) to handle multitasking and communications between software tasks. The design is naturally modular with a software bus allowing each task to communicate with other tasks as needed. Individual tasks are designed to handle each subsystem and perform all of its observing, telemetry and command tasks simultaneously.
Figure 3-48. Flight Software Module Interaction with Electronics and Software Bus
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During Phase A, the SoloHI flight software team developed and tested image processing algorithms on the RTEMS operating system using a LEON3 test chassis. This study evaluated the image compression time for baseline images to verify the LEON3 processing speed and validate the compression assumptions in the baseline SoloHI observing plan. The image compression times and factors for the same 1k x 1k linear gradient image using the Rice and H- Compress algorithms are presented in Table 3-28. The measured throughput rate using the LEON3 test chassis is 30.6 kbps/MIPS, assuming that the processing speed of the 25 MHz LEON3 FPGA is 20 MIPS. The required throughput rates for the Rice and H-Compress algorithms are ≥65.2 and ≥58.7 kbps/MIPS respectively. These throughput rates correspond to the SoloHI processing load estimate of 8.30 MIPS. The SoloHI team believes that there is an issue with the LEON3 implementation on the test chassis that does not allow it to achieve the published 20 MIPS. The SoloHI team will test performance metrics in early Phase B to determine the processing speed of the test chassis. Table 3-28. LEON3 FT Image Compression Summary
Algorithm Image Time (s) Compression Factor Rice 1K x 1K Linear Gradient 24 1.5-2.0 H-Compress (4) 1K x 1K Linear Gradient 24 3.5-4.5 3.2.1.2 Flight Software Requirements The SoloHI top-level flight software requirements are captured in Section 7 of the SoloHI Instrument Specification (NRL, SSD-SPC-SoloHI-0001) and in Section 4 of the SoloHI Instrument Requirements Document (NRL, SSD-RQT-SoloHI-0002). Section 7 of the SoloHI Instrument Specification defines the top-level requirements for Image Processing, Instrument Scheduling, Instrument Commanding, Instrument Housekeeping, and Instrument Autonomy Rules. Section 4 of the SoloHI Instrument Requirements Document defines the detailed image compression requirements for Image Processing. The SoloHI Flight Software Specification (NRL, SSD-SPC-SoloHI-0002) defines the architecture for the SoloHI flight software and the top-level requirements for each of the SoloHI flight software modules. 3.2.2 Ground Operations Software SoloHI will use the Integrated Test and Operations Software (ITOS) for command and telemetry handling. ITOS is a real-time control and monitoring system developed and maintained by a small team at GSFC. It is a portable, highly configurable system which runs under a variety of UNIX operating systems, including Solaris, FreeBSD, and Linux on workstation or PC hardware. It will be used at all stages of the SoloHI development, from initial FSW development and testing, verification and calibration and finally during operations. ITOS was used during SECCHI development and operations and therefore allows re-use of elements of the SECCHI ground system. 3.2.3 Software Test and Management Table 3-29 describes each build for the SoloHI flight software and ground software. Each software build is developed to satisfy a subset of the software requirements captured in the SoloHI Flight Software Specification (NRL, SSD-SPC-SoloHI-0002), the SoloHI Ground Software Specification (NRL, SSD-SPC-SoloHI-000X), and in the SoloHI Instrument
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Specification (NRL, SSD-SPC-SoloHI-0001). All SoloHI software requirements will be satisfied with the successful completion of the Build 4 build test. Table 3-29. Flight/Ground Software Build Strategy
Build Functionality Development Equipment 1 Core functionality and heritage SECCHI LEON3 FT test chassis, Spacewire USB - PC including: Spacewire drivers Basic Image Processing and Compression Timekeeping 2 Camera interface development, SCE DM, DRB DM and SoloHI Detector EM, Basic Image Scheduling Spacecraft Simulator for Single and Summed Images, Spacecraft Data Interfaces (Commanding, Science and Housekeeping) 3 Power control Electronics Development Models Spacecraft Power Interfaces (SCE DM, SPS DM, DRB DM, and SoloHI SoloHI Image Processing, Special Detector EM), Observations, Calibration, Spacecraft Simulator
4 Advanced Image Scheduling, Image Flight Spare Electronics Processing and Data Management, (SCE FS, SPS FS, DRB FS, and SoloHI Detector Housekeeping Collection EM), Flight Spare Harnesses, Flight Sensors,Spacecraft Simulator 4.x Bug Fixes Flight Spare Electronics (SCE FS, SPS FS, DRB FS, and SoloHI Detector EM), Spacecraft Simulator Each software build for both flight and ground software will be under version control starting with Build 1. Each software build will be tested by the SoloHI test engineer using automated scripts. Since the SoloHI software team is small, there will not be dedicated SoloHI test engineers. The SoloHI test engineer to test a specific software module for the given software build will be selected from the SoloHI software development team with the requirement that the developer of the software module cannot test his own code. Software metrics, code management and bug tracking reports will be submitted to the project on a regular basis. Critical and newly designed code will be subject to peer review. The development equipment for Software Build 1, including the LEON3 FT test chassis, was procured during Phase A and has been used to port the image processing algorithms to the RTEMS operating system. The LEON3 FT test chassis includes three Spacewire ports, so it can be used to fully develop the spacecraft primary and redundant Spacewire links which carry the housekeeping and science data to the spacecraft, and the commands to the SoloHI instrument from the spacecraft. NRL has tested the Spacewire ports on the LEON3 and verified operation. Software Build 1 testing shall be completed in 2012. Later builds incorporate higher fidelity hardware as it becomes available. Software Build 2 shall complete the spacecraft interface development and initiate the camera electronics interface development with the arrival of the SCE DM. Software Build 3 shall complete the camera
3-55 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 interface development by testing the complete camera system with the SCE DM, the SPS DM, the DRB DM, and the 4k x 4k SoloHI detector EM operating at the baseline 4 Mpixel/s readout rate over 2 ports simultaneously. Software Build 3 shall also complete the basic observing program modes and the calibration modes. Software Build 4 shall be developed on the SoloHI flight electronics spares (including the SCE FS, SPS FS, DRB FS and SoloHI detector EMs. Software Build 4 will develop the advance image scheduling and processing function and add a data manager to ensure that the complete 30-day mission timeline stays within its data volume. In addition, it will test the housekeeping data collection using the flight sensors. This software build will correct for any differences between the SoloHI electronics development models and flight models. Software Build 4 will be ported from the Flight Spare electronics to the Flight Model electronics, before the Instrument Pre-environmental Review Software Builds 4.1 to 4.x will be developed on the Flight Spare Electronics to address outstanding problem reports that are generated during the SoloHI environmental testing or after delivery to the spacecraft team for spacecraft integration. These software builds will not implement new features to satisfy the documented flight software requirements. The ground software will be used as early as possible to automate the test process, and the build strategy will parallel that of the FSW. The same ground software used for development will be used for I&T and flight operations. This will provide a library of ground procedures that can also be used in flight or adapted for flight.. 3.3 Technology Development Plan The SoloHI technology development plan is described in the SoloHI technology development plan document number SSD-PLN-SOLOHI-0001 and is available on request. 3.4 Spacecraft Interface and Accommodations 3.4.1 Interface Definition 3.4.1.1 Mechanical and Thermal Interfaces The SoloHI mechanical interfaces with the Solar Orbiter spacecraft are captured in Sections 4.3 through 4.6 of the SoloHI Experiment Interface Document, Part B (EID-B). The SoloHI thermal interfaces with the Solar Orbiter spacecraft are captured in Section 4.7 of the EID-B. 3.4.1.2 Electrical Interface The SoloHI electrical interfaces with the Solar Orbiter spacecraft are captured in Section 4.7 of the EID-B. The SoloHI thermal interfaces with the Solar Orbiter spacecraft are captured in Section 4.7 of the EID-B. 3.4.1.3 Command and Data Handling Interface The SoloHI command and data handling (C&DH) and software interfaces with the Solar Orbiter spacecraft are captured in Section 4.7 of the EID-B. The SoloHI thermal interfaces with the Solar Orbiter spacecraft are captured in Sections 4.9 and 4.10 of the EID-B. 3.4.2 Interface Requirements and Accommodations 3.4.2.1 Mass
3.4.2.1.1 SoloHI Allocation
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The mass allocation for the SoloHI science payload, including the SoloHI Instrument Module (SIM), including the SIM brackets, the SoloHI Power Supply box (SPS) and the wire harnesses, is 15.5 kg.
3.4.2.1.2 Current Estimate The SoloHI mass current best estimate (CBE) (also known as the SoloHI Basic mass) is 12.74 kg. The SoloHI mass CBE with contingency (also known as the SoloHI Nominal mass) is 14.57 kg, based on a mass contingency of 15.1%, or 1.92 kg. The mass contingency percentage for each SoloHI subsystem is based on the maturity of the basis of mass estimate (BME). Table J-1 defines the basis of mass estimate codes. The Solar Orbiter Project Office at GSFC, in consultation with the Heliophysics Division, has defined the mass margin for the Solar Orbiter instruments to be the difference between the SoloHI mass allocation and the mass CBE. This mass margin definition is different than the margin definition in the NASA Gold Rules, Revision E that is based on the difference between the mass allocation and the mass CBE with contingency. The NASA Gold Rules requires that instruments shall have a mass margin of ≥25% at the end of Phase A. The Solar Orbiter Project Office at GSFC has directed the U.S. Solar Orbiter instruments to compare the mass margin based on the difference between the allocation and CBE with this requirement. The SoloHI team will continue to report the SoloHI nominal mass (CBE + contingency), to assess the risk of exceeding the SoloHI allocation. The SoloHI mass margin between ths mass allocation and CBE is 2.76 kg (or 21.6%). The difference between the SoloHI mass allocation and CBE with contingency is 0.93 kg (or 6.4%).
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Table 3-30 provides the mass breakdown for the SoloHI science payload.
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Table 3-30. Mass Breakdown for the SoloHI Science Payload
The current SoloHI mass margin exceeds the Gold Rules mass margin requirement of 25% at the end of Phase A, but remains below the mass margin requirement of 20% at the end of Phase B. Since The SoloHI project has just completed a year-long Extended Phase A based on the NASA decision to rescope the SoloHI design to reduce costs. In this period, the SoloHI instrument redesign has progressed to a higher design maturity than typically reached in the Phase A period. Therefore, the SoloHI team is not concerned about this Gold Rules mass margin violation, because the SoloHI design is still on track to satisfy the mass margin requirement at the end of Phase B.
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3.4.2.1.3 Changes from the December 2010 CSR Table 3-31 shows the difference Table 3-31. SoloHI Science Payload Mass Change since the 12/10 CSR between the current SoloHI mass CBE with contingency and the mass CBE w/ contingency captured in the December 2010 CSR. The SoloHI mass CBE with contingency has increased from 12.45 kg to 14.57 kg for a total mass delta of 2.13 kg since the release of the December 2010 CSR. The primary reasons for the mass growth includes the addition of the SIM brackets to the SoloHI science payload (2.90 kg), the reduction in the SPS to SIM/SCE wire harness by moving the SoloHI electronics to the spacecraft deck exterior under the SIM (-1.98 kg), the increase in the SIM brackets, the SIM structure, and SCE enclosure to support the increased electronics mass with the combined SoloHI control and camera electronics (0.50 kg), the addition of the Thermal Isolation Package in the SIM focal plane assembly to thermally isolate the detector readout electronics from the detector (0.18 kg), the increase in the thermal components to maintain the standard 3% of total mass rule of thumb (0.03 kg), increased maturity in the SCE and interior baffle designs (0.497 kg). ESA increased the SoloHI mass allocation by 1.0 kg for the SIM bracket addition to the SoloHI science payload after accounting for the shift in the SoloHI electronics location from the spacecraft interior to the +Y spacecraft panel exterior under the SIM that reduced the length and mass of the SoloHI inter- assembly wire harness. 3.4.2.2 Operational Average Power
3.4.2.2.1 SoloHI Allocation
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The operational average power allocation for the SoloHI science payload, including the SoloHI camera electronics, the SoloHI power supply (SPS) box and the operational heaters, is 12.0 W.
3.4.2.2.2 Current Estimate The SoloHI operational average power current best estimate (CBE) (also known as the SoloHI Basic power) is 8.95 W for the Perihelion observation period and 9.44 W for the Out-of-Ecliptic observation periods. The SoloHI operational average power CBE with contingency (also known as the SoloHI Nominal power) is 10.29 W for the Perihelion observation period and 10.85 W for the Out-of-Ecliptic observation periods, based on a contingency of 15%. The contingency is 1.34 W and 1.42 W for the Perihelion and Out-of-Ecliptic observation periods respectively. The average power contingency percentage for each SoloHI subsystem is based on the maturity of the basis of power estimate (BPE). The Solar Orbiter Project Office at GSFC has defined the average power margin as the difference between the SoloHI power allocation and the power CBE. This definition is different than the definition in the NASA Gold Rules, Revision E, that is based on the difference between the average power allocation and the average power CBE with contingency. The NASA Gold Rules requires that instruments shall have an average power margin of ≥25% at the end of Phase A. The Solar Orbiter Project Office at GSFC has directed the U.S. Solar Orbiter instruments to compare the power margin based on the difference between the allocation and CBE with this requirement. The current SoloHI average power margin is 2.56 W (or 27.1%), based on the margin between the SoloHI average power allocation and the average power CBE. The current SoloHI power margin satisfies the 25% Gold Rules average power margin requirement for Phase A. The difference between the SoloHI average power allocation and CBE w/contingency is 1.34 W (or 14.2%).
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Table 3-32 shows the average power breakdown for the SoloHI science payload.
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Table 3-32. Average Power Breakdown for the SoloHI Science Payload
3.4.2.2.3 Changes from the December 2010 CSR Table 3-33 shows the difference between the current SoloHI average power CBE with contingency and the average power CBE w/ contingency captured in the December 2010 CSR. The SoloHI average power CBE w/ contingency has decreased from 11.01 W to 10.85 W for a total average power delta of -0.16 W since the release of the December 2010 CSR. The average power reduction includes the consolidation of the control electronics and the A/D converters into the SCE (-0.49 W), and the increased operational heater power to maintain the lens barrel, detector, and operational amplifier temperatures within the cold operational limits (0.33 W). The SoloHI average power allocation of 12.0 W has not changed since the release of the December 2010 CSR.
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Table 3-33. SoloHI Science Payload Operational Average Power Change since the December 2010 CSR
3.4.2.3 Envelope
3.4.2.3.1 SoloHI Allocation The SoloHI Instrument Module (SIM) static envelope in its launch configuration (with the door closed) is 63.0 cm XOpt x 21.0 cm ZOpt x 42.0 cm YOpt. The SoloHI Power Supply (SPS) box envelope is 16.85 cm XOpt x 17.0 cm ZOpt x 23.0 cm YOpt. This envelope includes two keepout envelopes of 5.5 cm XOpt x 20.15 cm ZOpt x 7.82 cm YOpt on the spacecraft interface connector panel and instrument interface connector panel sides.
3.4.2.3.2 Current Estimate
The current SIM dimensions (with the door closed) are 61.7 cm XOpt x 19.8 cm ZOpt x 40.25 YOpt. The SIM envelope margin is 1.3 cm XOpt, 1.2 cm ZOpt, and 1.75 cm YOpt. The current SPS dimensions are 16.70 cm XOpt x 14.36 cm ZOpt x 17.07 cm YOpt. The SPS envelope margin on each side is 0.15 cm XOpt, 2.64 cm ZOpt, and 0.63 cm YOpt.
3.4.2.3.3 Changes from the December 2010 CSR The SIM length increased from 58.5 cm to 61.7 cm in length and from 18.6 cm to 19.8 cm in width compared to the design in the December 2010 CSR, due to the placement of the SCE on
3-64 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 the aft face of the SIM FPA. The SIM width increased from 18.6 cm to 19.8 cm compared to the design in the December 2010 CSR, due to typical design maturity. The SIM height increased from 22.8 cm to 40.25 cm, because the responsibility for the SIM brackets under the SIM structure were transferred from Astrium to NRL during the Extended Phase A period. The current SPS dimensions are smaller than the SoloHI Control Electronics dimensions in the December 2010 CSR. The SPS length has reduced from 20.91 cm to 16.70 cm, while the width has reduced from 17.61 cm to 14.36 cm. The SPS card have a smaller board area and there are only 2 cards in the SPS enclosure compared to the 3 cards in the SoloHI Control Electronics enclosure. The SPS height has increased from 6.55 cm to 17.07 cm, because the SPS cards are vertically oriented compared to the horizontal orientation for the cards in the SoloHI Control Electronics. 3.4.3 Pointing Requirements The SoloHI pointing accuracy, pointing jitter and pointing windowed stability requirements for the SoloHI instrument during SoloHI science operations are captured in Section 3 of the SoloHI Instrument Requirements Document. The corresponding pointing requirements for the Solar Orbiter spacecraft during SoloHI science operations are captured in Section 4.4.5 of the SoloHI Experiment Interface Document, Part B (EID-B). The spacecraft pointing metrics include the absolute pointing error (APE), the relative pointing error (RPE) and the pointing drift error (PDE) and are defined in Section 4.4.2.2 of the SoloHI EID-B. 3.4.4 Instrument Alignment Requirements
The nominal location of the SoloHI F1 baffle tip relative to the trailing edge (-XOpt) of the heat shield sun shade shall be located at [- 68.27 cm (XOpt), -3.45 cm (YOpt)], as shown in Figure 3-49. The SIM mounting interface is 17.16 cm above the spacecraft deck with the current heat shield dimensions. The SIM instrument mounts will have a shimming capability to adjust the SIM height and distance from heat shield to satisfy the required heat shield alignment tolerance during the observatory integration and test activities. The alignment tolerance of the SIM F1 baffle tip with the trailing edge of the Solar Orbiter heat shield sun shade on-orbit during science observation shall be ±20 mm (XOpt) and ±9.2 mm (YOpt). This tolerance shall apply to a 22 cm section of the heat shield leading edge that is centered where the SoloHI XHIO-ZHIO plane intersects with the heat shield sun shade. This tolerance would permit the SoloHI instrument to capture the 5° inner FOV at 0.28 AU and satisfy the SoloHI scene coverage requirement. The SIM will satisfy the SoloHI straylight rejection performance requirements at the worst-case Solar Orbiter off-pointing of ±1.0° based on the 0.28 AU perihelion using the worst-case F1 baffle to heat shield alignment tolerance. The SoloHI alignment tolerance budget is derived in Appendix B of the SoloHI Instrument Specification. The alignment interface between the spacecraft and the SIM is captured in the SoloHI EID-B.
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Figure 3-49. SIM Front Baffle to Heat Shield Leading Edge Nominal Location and Alignment Tolerance 3.4.5 Instrument Field of View The SoloHI instrument field of view (FOV) extends across 40° radially, and a maximum of 40° tangentially. However, since the inner FOV is vignetted by the SoloHI forward baffles, the SoloHI FOV must be specified with two separate FOV pyramids, each with its own vertex. The vertex of the unvignetted inner FOV pyramid in the SoloHI optical reference frame is located at [-44.715 cm XHIO, 0.0 cm YHIO, -5.356 cm ZHIO]. The unvignetted inner FOV is 33.425° radial and 40.0° tangential, with a boresight that is oriented 24.850° from the Sun geometric center. The vertex of the vignetted FOV pyramid in the SoloHI optical reference frame is located at the top of the A1 aperture at [-42.636 cm XHIO, 0.0 cm YHIO, -3.512 cm ZHIO]. The vignetted FOV is 3.208° radial and 33.940° tangential, with a boresight that is oriented 6.536° from the Sun geometric center. 3.4.6 Instrument Unobstructed Field of View The SoloHI instrument unobstructed FOV (UFOV) is defined as azimuth-elevation angles with respect to the F1 front baffle, the peripheral baffles, and the AE2 baffle on the light trap. The peripheral baffle UFOV is defined at 3 vertices: the AE2_LV2 vertex and the I1_LV1 vertices at the aft and front of the PL1 peripheral baffle respectively, and the F1-LV1 at the front of the PL2 peripheral baffle. A symmetric set of 3 vertices are used to define the UFOV on the PR1 and PR2 peripheral baffles. Table 3-34 defines the SIM UFOV boundary as a series of discrete lines defined by its base vertex location and whose orientation is described by the azimuth/elevation angle pair. The base vertex location is defined in the SoloHI optical reference frame (HIO). The azimuth-elevation coordinate transformation is a positive rotation about the ZHIO axis for the given azimuth angle, followed by a negative rotation about the YHIO axis for the given elevation angle. Figure 3-50 shows the SoloHI UFOV elevation as a function of the azimuth. No surface on the spacecraft
3-66 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 shall extend above the SoloHI UFOV defined in Table 3-34. In particular, the Solar Orbiter heat shield, the solar arrays on the +YOpt panel, and the RPW antenna on the +YOpt panel shall not extend within the SoloHI UFOV. 3.4.7 On-Orbit Calibrations The SoloHI science payload shall perform a photometric calibration sequence at least twice each orbit before and after the 30 day observation period for each orbit. This calibration shall require SoloHI measurements of a visible light calibration star. Based on the wide FOV of the SoloHI instrument, no spacecraft maneuver is required to capture a standard set of calibration stars, similar to those used for the LASCO and SECCHI on-orbit calibrations. The purpose of the SoloHI photometric calibration is to measure the SIM telescope throughput loss during the mission. The final photometric calibration using the standard stars is accurate to ~3%, which is well within the photometric accuracy requirements of 5% at the inner FOV and 20% for the outer FOV of SoloHI Full Frame images. Table 3-34. SIM Unobstructed FOV Definition
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Figure 3-50. SIM Unobstructed FOV Boundary Elevation Over the Azimuth Range 3.4.8 Cleanliness Requirements The SoloHI Contamination Control Requirements and Plan (CCRP) (NRL, SSD-RQT-SoloHI- 0004) defines the SoloHI surface cleanliness levels (particulate and molecular) for internal and exterior surfaces at the baseline mission EOL. The SoloHI CCRP derives the SoloHI surface cleanliness budget throughout the SoloHI AIT flow including subsystem delivery for science payload integration, instrument delivery for spacecraft integration, launch and the end of the baseline science mission. 3.5 Manufacturing, Integration, and Test 3.5.1 Manufacturing, Integration, and Test for SoloHI Units and Assemblies The SoloHI science payload includes the SoloHI Instrument Module (SIM), the SoloHI Control Electronics box (SCE) and the SIM to SCE wire harnesses. The following section describes the manufacturing, integration and test program for the SoloHI units and assemblies before they are delivered to NRL for the SIM integration and the overall SoloHI science payload integration. The integration and test program for the SIM and SoloHI science payload is discussed in Section 3.5.2. The organization that is responsible for the manufacture, integration and test of each of the SoloHI units and assemblies before SIM or SoloHI science payload integration is defined in Table 3-35. NRL is responsible for the assembly and test of the Lens Barrel Assembly, Focal Plane Assembly, SCE, camera electronics, and Baffle Cover Door Assembly. ATK will assemble and test the SIM structure, instrument mounts, straylight baffle assemblies, and FPA structural elements. Sarnoff Corporation will assemble and test the APS detectors.
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All SoloHI components and parts are fabricated using standard, proven manufacturing processes for space flight hardware and adhere to the mission assurance requirements of the SoloHI PAIP. The manufacturing operations are controlled using Work Order Authorizations (WOAs) that authorize work, divide the overall work into discrete tasks, provide a log for manufacturing events and testing, and is used for reporting problems, failures, and anomalies. All assembly, integration and test (AIT) plans follow the configuration control, mission assurance and contamination control requirements of the SoloHI program. Table 3-35. Providers for SoloHI Units and Assemblies Assembly Provider Structure ATK Beltsville Lens Barrel Assembly Qualified vendor Forward and Interior Baffle Assemblies ATK Beltsville Active Pixel Sensor Sarnoff Corp Camera Electronics (Detector Readout Board, Camera interface Electronics) NRL Code 8200 Baffle Cover Door Assembly NRL Code 8200 SoloHI Control Electronics NRL Code 8200 3.5.1.1 SIM Structure The SIM rigid box structure is comprised primarily of CFRP panels, using a M55J lay-up with cross orientation. The structure fabrication of the CFRP structure is straight forward, and will draw on flight proven, industry standard processes for the panels, similar to those used for the SCIP assembly closeout structure and in the SECCHI/HI structure. The primary structure is composite lay-up of CFRP with cross orientation. As in the SECCHI/HI assembly, doublers of similar material will be used locally to stiffen the structure at the interface of the instrument mounts and door and associated hardware attachment. The FPA assembly is mounted within the primary structure via a Ti-6AL-4V housing. The FPA assembly is mounted to the primary structure via a Ti-6AL-4V housing. The CFRP structure walls and door assembly are joined together using standard, qualified techniques for box enclosures. Panel clips and doublers will be used at the seams. The forward baffles are attached to the ledge via flexure clips and alignment provisions are incorporated for anticipated adjustments in the baffle tips during alignment testing. The internal baffle box is a self contained subassembly housing the interior baffles which will be treated with an optical dark black coating. This manufacturing and assembly approach has been tailored to minimize the handling of the interior baffles. The baffle box side walls have slots and bosses to accept attachment of the interior baffle tabs external to the box. Precision machining of the side wall bosses will provide self alignment of the baffle tips. Final alignment of the baffles system during integration will be accomplished with shims. The final door positioning and adjustments are made during mechanism installation and testing. NRL will construct and test a structural model and a flight model of the SIM. The structural model will be used for structural design validation activities and will also support the ESA spacecraft level testing. The flight model structure will undergo protoflight testing at the integrated instrument level.
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3.5.1.2 Lens Barrel Assembly The SoloHI lens barrel assembly will be fabricated and tested by a qualified vendor under NRL supervision. During Phase B, NRL will select the vendor from a group of candidate optical vendors that NRL has used in the past. The vendor will complete the general design by the SoloHI PDR and present final design at a Lens Barrel Assembly CDR before the overall SoloHI CDR. After this review, NRL will approve the fabrication of three lens assemblies. The vendor will provide materials, fabrication, assembly and test documentation with the lenses. The lenses will undergo optical performance verification and characterization at NRL. One of the lens assemblies will undergo qualification testing. A second lens will be selected as the flight unit. 3.5.1.3 Straylight Baffle Assemblies The straylight baffle development model and flight model assemblies will be provided by ATK Beltsville. The mechanical design will be completed at ATK Beltsville to an optical prescription provided by NRL. The aluminum baffle parts will be procured from ATK qualified vendors. The design includes provisions for shimming critical dimensional tolerances; the remaining tolerances are maintained using precision machining. Shimming is introduced where required to meet the tolerances. An optical dark black coating will be applied to the baffle surfaces. The baseline selection for the optical dark black coating is Epner laser black. A qualification model baffle assembly will be fabricated and tested. Thisunit will allow complete end to end testing of the baffle design. An early design of the baffle qualification model, shown in Figure 3-51, has been completed and quotes for piece part fabrication have been received by local machine shops during Phase A. This design will be refined early in Phase B to satisfy all qualification model requirements and the qualification model procurement will be initiated after the Baffle Assembly peer review in Phase B. The back end of the mock up unit accommodates a test camera to facilitate system level testing. Baffle coupons for deep space black and Epner laser black were procured during Phase A. This procurement investigated options for edge shapes and minimal baffle thicknesses and provided an example of the expected delivery times. NRL has conducted optical testing using coupons of the deep space black coating and the Epner laser black to compare with expected performance. The detailed design of the baffle assembly qualification model will allow the design team to solve several significant assembly and alignment issues. The optical qualification model will be essential for the stray light test program at NRL and then at CSL. The flight model related optical testing is described in Section 3.1.9.
Figure 3-51. Development Model Baffle Assembly
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3.5.1.4 APS Detector The flight APS detector will be provided by Sarnoff Corporation. Sarnoff Corporation has extensive experience providing detectors for high reliability space and terrestrial applications. The most recent flight program was the detector for the day-night instrument camera on NPOESS. Processes and procedures similar to those developed to package and test the NPOESS detector will be followed for SoloHI. The APS design is an adaptation of the Minimal APS design demonstrated in Sandboxes one through five. The principal adaptations of this design for SoloHI are the full well capacity and the pixel size. A test device meeting the SoloHI device requirements is under fabrication and will be tested during Phase B. Sarnoff Corporation has developed on-chip APS circuitry to provide a double correlated sample and hold function. The Minimal design architecture uses an off-chip ADC. SoloHI intends to use the latest version of this on-chip circuitry verified in the Sandbox VI Jazz run. The Sarnoff Corporation is responsible for the design, fabrication, assembly and test of the APS detector. Significant Sarnoff industrial partners include Chronicle Corporation and Jazz Semiconductor (owned by Tower Semiconductor Corporation). Chronicle Corporation will adapt the present design into a stitchable design and will design the non-imaging portions of the detector. Jazz Semiconductor will fabricate the device. All wafer screening testing, device packaging activities and flight APS screening operations will take place at the Sarnoff facilities in Princeton, NJ. APS screening procedures will include temperature cycling and high temperature burn-in. Selected screened devices will undergo qualification burn-in and radiation testing. 3.5.1.5 SoloHI Camera Electronics The SoloHI camera electronics will be provided by Code 8000 of NRL. Under NRL direction, Sarnoff Corporation, Silver Engineering and Orbital Sciences Corporation will make significant contributions to the PRC, CEC and DRB boards. Sarnoff Corporation will provide the conceptual design of the circuits which directly interface to the flight device on the detector readout board. Silver Engineering is responsible for the card design, conducting board level testing, and supporting the assembly testing. Orbital Sciences Corporation will provide parts engineering and board fabrication/population support. Code 8000 will provide electrical systems engineering, and conduct integrated electrical testing of the APS, DRB, and SCE system. The SoloHI team will build an engineering model DRB and SCE. The performance of this set of electronics will be tested with the flight and engineering detectors. Similar processes will be used to construct the flight DRB and the flight SCE. The flight SCE will undergo vibration and thermal vacuum testing after integration with the SoloHI flight telescope and prior to the full SoloHI instrument integration. 3.5.1.6 Focal Plane Assembly (FPA) The focal plane assembly includes the FPA housing, Detector Readout Board (DRB), detector Kovar mount with APS detector, DRB support frame, cold finger, and radiator. The development and manufacturing of all these FPA subassemblies, except for the DRB and detector Kovar mount with APS detector, will be conducted by ATK Beltsville under the management of the NRL SoloHI team. The ATK design group is the same group that developed the LASCO and SECCHI FPAs. ATK will procure the SoloHI FPA parts from qualified vendors and fabricate the FPA structural elements using certified equipment for all metal machining operations. The SoloHI FPA design builds on the “lessons learned” during the SECCHI FPA program in the
3-71 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 areas of surface treatments. The DRB will be provided by the NRL Code 8100 based on the Sarnoff design. The Kovar mount with APS detector will be provided by Sarnoff. A qualification model FPA with a flight-like APS and DRB will be used to verify the FPA design. Finite element modeling and thermal modeling will be used in conjunction with qualification testing and thermal testing. The flight model FPA will be assembled and delivered to instrument integration. 3.5.1.7 Baffle Cover Door Assembly The SoloHI baffle cover door assembly includes the door structural lid, the door hinges, the door torsional spring, and the door deployment mechanism. The overall baffle cover door assembly will be designed, assembled and tested by the NRL Code 8000 mechanism group led by Steve Koss. The door deployment mechanism is similar to the SECCHI/HI door and utilizes a HOP actuator, provided by the Starsys group at SpaceDev. The design is described in Section 3.1.8 of this report. The door components (hinges, latches, etc.) will be supplied by qualified vendors. The SoloHI program includes a life cycle model door and a flight model door. 3.5.1.8 SoloHI Power Supply The SoloHI Power Supply (SPS) boards are fabricated using standard processing techniques for space flight electronics. NRL reviews all fabrication plans and practices to monitor adherence to SoloHI requirements from the PAIP, CCP and CM. NRL uses a qualified space vendor, Orbital Sciences Corporation (OSC) for fabrication of each board assembly, which is then tested by NRL prior to integration into any sub-assembly with flight components. The same approach is used for the SCE elements. The SPS baseline package stacks the Power Control Card (PCC) and the Power Converter Card (PCNV) in a two card enclosure as shown in Figure 3-24. The SPS is mounted on the exterior of the Solar Orbiter +Yopt spacecraft panel under the SIM. Internal interconnect between the processor and power supply card is accomplished with a flight quality motherboard and daughter board style connection. Following standard practice at NRL for heritage electrical design, Code 8000 will build and test a SPS development model (DM) an SPS structural thermal model (STM) and a SPS protoflight model. The SoloHI team will The SPS FM will undergo vibration, thermal vacuum testing and preliminary EMI/EMC testing prior to integration with the SoloHI instrument. An SPS structural thermal model will be provided to ESA to support spacecraft level mechanical and thermal testing. we will deliver the EDM to ESA to support spacecraft level engineering model electrical testing. 3.5.2 Integration, Test, and Verification The following sections provide an overview of the SoloHI science payload integration, the verification and validation approach, and the verification test description. In addition, the SoloHI test programs for environmental, functional, performance, characterization and variance tests are described. Figure 3-52 describes the entire SoloHI assembly, integration and test (AIT) flow before delivery for spacecraft integration. It includes the initial integration of the SoloHI Instrument Module, integration of the SoloHI science payload, the environmental test program, and instrument delivery for spacecraft integration. Functional, performance and characterization tests are
3-72 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 performed before and after the entire environmental test sequence. The SoloHI test processes, failure and problem reporting, and documentation shall comply with the SoloHI Product Assurance Implementation Plan and the SoloHI Configuration Management Plan. These top- level SoloHI documents are fully compliant with the applicable NASA test and verification requirements.
Figure 3-52. SoloHI Assembly, Integration, and Test Flow 3.5.2.1 Science Payload Integration The assembly of the SoloHI Instrument Module (SIM) will integrate the SIM telescope, SoloHI baffle assemblies, SoloHI baffle cover door assembly, and Camera Interface Electronics (CIE) box, CIE to FPA wire harness, thermal components and instrument mounts on the SIM structure. The SIM will be integrated in the clean room facilities at NRL.
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3.5.2.2 Verification and Validation Approach The SoloHI Verification and Validation Plan describes the SoloHI verification approach and documentation. The SoloHI verification matrix will track the verification status of all SoloHI requirements from Level 1 Project requirements to Level 4 instrument specification and interface requirement documents. The SoloHI team shall develop a SoloHI verification matrix that shall identify the verification method, the hardware level of assembly for each verification, the parent requirement for each SoloHI requirement, and the verification documentation. By the SoloHI PDR, the preliminary SoloHI verification matrix shall be prepared that will define the verification method, the verification method description, and the requirement traceability using the parent/child relationships for all SoloHI Level 1 to Level 4 requirements. At the SoloHI PSR, the final verification matrix will be submitted to the NASA Solar Orbiter Project Office as part of the SoloHI Acceptance Data Package. The SoloHI verification matrix will be available for review at the SoloHI PER and shall be available to review on request by the NASA Solar Orbiter Program Office representatives at NRL after the SoloHI CDR. The following sections provide definitions for the SoloHI requirement levels, hardware levels of assembly, and hardware models.
3.5.2.2.1 Requirement Level The SoloHI project requirements in each requirements document are classified at a specified requirement level: Level 1 requirements are the Project‟s fundamental set of top-level requirement levied by the Program or captured at the Program level. Both NASA Headquarters and ESA will have Level 1 requirements for the Solar Orbiter project. The Level 1 requirements shall define the scope of the science objectives, describe the science measurements to achieve those objectives and define the success criteria for the Baseline science mission and Threshold science mission. The Level 1 requirements shall also include the Program Management, Systems Engineering, and Verification requirements levied by NASA Headquarters on all NASA projects. Level 2 requirements are the Project‟s fundamental set of top-level requirement including the mission requirements defined by ESTEC and the program requirements defined by the NASA Project Office. The Level 2 requirements shall describe the mission segments and set the top-level requirements on the spacecraft bus, launch vehicle, and the ground segment. The mission timeline and the resource allocations for the combined science payload shall be also be described by these requirements. The Level 2 requirements shall include the Mission Assurance requirements levied by the Project. Level 3 requirements are the top-level requirements for the science payloads. The Level 3 requirements shall include the detailed instrument science requirements derivation for each of the Level 1 science objectives/questions that results in a detailed instrument science measurement requirements for the Baseline and Threshold science missions. The Level 3 requirements will capture the top-level instrument performance, environmental and accommodations requirements. The Level 3 requirements will also include the Instrument Project plans for Project Management, Systems Engineering and Product Assurance.
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Level 4 requirements include spacecraft, science payload and mission operations specifications, and interface control documents between the spacecraft and the instrument, and between the observatory mission operations center (MOC) and the instrument payload operations center (POC). Level 5 requirements include instrument subsystem specifications for all hardware and software subsystems to be designed or procured, including electronics, detectors, mechanisms, optics, structures, flight software, and ground systems. The SoloHI requirement flowdown from the Level 1 Program requirements to the Level 3 SoloHI Science Payload requirements is shown in Figure 3-53, while the SoloHI requirements flowdown from the Level 3 SoloHI science requirements to the Level 5 SoloHI subsystem specifications is shown in Figure 3-54. A specific requirement document may contain requirements at 2 different requirement levels.
Figure 3-53. SoloHI Requirements Flowdown from Level 1 to Level 3
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Figure 3-54. SoloHI Requirements Flowdown from Level 3 to Level 5
3.5.2.2.2 Hardware Levels of Assembly Figure 3-55 defines the hardware levels of assembly and their hierarchy from highest level of assembly to lowest. The Observatory is defined as the Spacecraft Bus integrated with the Science Payload. The Science Payload is a group of payload modules, units, and assemblies that comprise the entire instrument experiment. The Module is a major subdivision of the instrument payload consisting of multiple units and assemblies that are physically integrated. The Subsystem is a functional subdivision of the Spacecraft Bus or Science Payload that include a group of units, assemblies and components that are not physically integrated.
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Examples include the Spacecraft Structural, Attitude Determination and Control, Command and Data Handling, Power, and RF Communications subsystems. The Unit is a functional subdivision of a module that is a self-contained physical unit which can perform a function that is required for the operation of the overall Assembly. Examples of units can include instruments, structures, and electronics boxes. The Assembly is a physical subdivision of the unit and is a group of subassemblies, components and parts. Examples include instrument integrated optical lens assembly, baffle assemblies, and focal plane assemblies that are integrated to achieve tight alignment requirements. The Component (also known as the Subassembly) is a functional subdivision of a Unit consisting of multiple parts which performs a function that is required for the operation of the overall Unit. Examples of components can include instrument mechanisms, kinematic or flexure mounts, a circuit card in an electronics box, or wire harnesses. The Part is a hardware element that cannot be disassembled or subdivided further without destroying its designed use. Examples of parts can include fasteners, connectors, and electronic parts such as resistors, FPGAs, memory chips, and the APS detectors. A subsystem of the Instrument Payload is defined as a logical subdivision of the Instrument Payload higher than the Part. The subsystem classification may be used to refer to Assemblies, Units or Components. For example, the SoloHI requirement and verification documents will refer to the SIM, the SIM and SCE structures, the SoloHI electronics boxes, and SoloHI baffle cover door mechanism as subsystems. Figure 3-56 presents the SoloHI levels of assembly from the Observatory to the Part level.
Figure 3-55. Hardware Levels of Assembly Hierarchy
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Figure 3-56. SoloHI Hardware Assembly Levels (Expanded to Unit Level)
3.5.2.2.3 Hardware Models The SoloHI hardware models are defined based on their intended use and their hardware level of assembly. The Breadboard Model (BB) is a hardware part or component that is built early in the design phase to verify that the design or fabrication process will meet the functional or performance requirements during and following exposure to a similar environment that the subsystem will experience during the mission.
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The Development Model (DM), also known as the Engineering Model (EM) is a hardware component or unit, that is built before the flight hardware design is placed under configuration control, to reduce performance, cost or schedule risk associated with critical hardware items or hardware items with little design margin. The development model should be similar to the flight design, in terms of the mechanical, optomechanical, electrical, or optical properties, that will be verified by the test. In addition, the development model should be fabricated and assembled similar to the flight model, if possible. The Life Cycle Model (LM) is a hardware component that will be used to verify that a mechanism can operate for the number of cycles that it will experience over its lifetime. The mechanism lifetime shall include the number of cycles that it performs during ground testing, instrument on-orbit precommissioning and the spacecraft mission. The Qualification Model (QM) is a hardware component, unit or assembly that will be used to qualify the flight hardware design for the launch and space environments. The qualification model will not be used as the flight hardware, unless it is disassembled, visually inspected, and refurbished with new parts or components as needed. The Flight Model (FM) is the hardware at all levels of assembly that is intended to be used for the actual mission. The Flight Spare Model (FS) is the hardware at all levels of assembly that may be used for the actual mission to replace a Flight Model which is no longer acceptable for flight. A formal System Verification Plan (Section 3.5.2.3) defines in detail tests required to demonstrate acceptability of the H/W and S/W configurations. Prescribed tests are completed using approved procedures, and documented in a final report. These tests provide design qualification, and verify workmanship, material integrity, and readiness for flight. The verification test program (Section 3.5.2.3) includes FSW validation to ensure that all S/W requirements are met. 3.5.2.3 Verification Test Description The SoloHI verification test program will include environmental, functional, performance and variance tests. Characterization tests do not verify subsystem requirements, but they can be used to estimate system model parameters and thereby validate subsystem models using test measurements.
3.5.2.3.1 Environmental Tests Environmental tests verify the hardware design, fabrication process, and hardware workmanship for individual units. Hardware qualification tests are used to demonstrate that the hardware design and its related manufacturing process will meet the subsystem design requirements. Acceptance tests and inspections are used to demonstrate that flight hardware is free of defects and assembly errors and is ready for mission use. Acceptance test methods shall be designed to find failures due to infant mortality and workmanship at each level of assembly.
3.5.2.3.2 Functional Tests Functional tests are performed to verify the basic functionality of electronics, mechanisms, sensors, actuators, heaters, flight software and ground software. Functional tests can be divided
3-79 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 into two categories: abbreviated functional tests and full functional tests. Abbreviated functional tests, also known as aliveness tests, will verify that electronics and mechanisms are still operational throughout the I&T program using the existing electronics architecture. Abbreviated functional tests will not test every mechanism position or operation and will not test any electronics or mechanism function that requires a special test configuration with external ground support equipment. The abbreviated functional test is intended to be short (usually less than 1 hour) and not require much time for the test setup, so that it can be performed often throughout the I&T program. The subsystem full functional test will verify that the electronics and mechanism will satisfy all functional requirements. Electronics functional test must exercise all electronics power modes, different states for the electronics (switch positions, etc.), communications paths between electronics boxes, cards, mechanisms, and sensors, and different processing loads. The flight software functional tests shall exercise all software modes, software mode transitions, and every software routine. Mechanism functional test shall verify that the mechanism can move to each of their positions and test each type of motion between positions. Door functional test shall verify that the door can move to each of its positions and deploy successfully from any closed position in the intended environment. Sensor functional test will show that the sensor measurements are still valid, while actuator functional tests verify that the actuators can move over the entire range of motion and to the commanded position. Heater functional tests will verify that the temperature is repeatable for the input voltage. The full functional test may require special test configurations with external ground support equipment.
3.5.2.3.3 Performance Tests Performance tests are performed to verify the performance requirements for instruments, structures, electronics, mechanisms, and control systems. Performance requirements are written such that the test measurement must be compared to the requirement value. This requirement can be written that the test measurement is higher than a threshold value, lower than a ceiling value, or within an acceptable range. In certain cases, the performance requirement can be written as a bounding one-dimensional function (y=f(x)). The performance test measurement and the added measurement uncertainty (or 99% confidence interval) should satisfy the performance requirement. Performance tests typically required external GSE and special test facilities to take these measurements.
3.5.2.3.4 Characterization Tests Characterization tests are performed to obtain test measurements to estimate parameters in system models, or to provide a complex baseline measurement of subsystem performance. Characterization tests are not performed to verify system functional or performance requirements. Variance tests can measure the change in the characterization test measurements from the baseline characterization. This will check qualitatively whether the system has been altered since the baseline measurement.
3.5.2.3.5 Variance Tests Variance tests measure the change in the instrument state or performance from the initial baseline measurement, and determine if the measurement change from transportation and handling, testing or exposure during the environmental test program, satisfies the related
3-80 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 performance requirement. Many of the instrument requirements, including alignment, spatial resolution, optical throughput, and particulate and non-volatile residue surface cleanliness, include derived requirements for the allowed performance degradation from the initial verification after instrument assembly to the last verification check before launch. 3.5.2.4 Environmental Test Program The SoloHI environmental test program is captured in the SoloHI Verification and Validation Plan. It is designed to satisfy the NASA and ESA Project-level requirements captured in the GSFC General Environmental Verification Standard (GEVS), the GSFC Rules for the Design, Development, Verification and Operation of Flight Systems (NASA Gold Rules), and the Solar Orbiter Experiment Interface Document, Part A (EID-A). The SoloHI environmental test program is presented in the following sections, divided by the hardware levels of assembly at the Unit level or higher and those at the Assembly/Subassembly levels. Table 3-36 shows the SoloHI environmental test matrix for SoloHI hardware levels of assembly from Subassembly level to the Observatory level. The table is divided between the SoloHI subsystem Flight Models and the Qualification/Risk Reduction Models.
3.5.2.4.1 SoloHI Flight Model The SoloHI science payload is comprised of the SoloHI Instrument Module (SIM), the SoloHI Power Supply (SPS) box, and the SoloHI wire harnesses connecting these two physical units. The SIM and the SPS will both be exposed to protoflight levels and durations in separate vibration tests. The SIM will undergo both random and sine vibration tests, while the SPS will only undergo a random vibration test. The overall SoloHI science payload will undergo thermal vacuum cycling, thermal balance, and EMI/EMC testing. Table 3-36 shows the number of operational and survival cycles and the minimum number of hours for the thermal soak duration in the thermal vacuum cycling test. The number of door deployments at the Cold Door Deployment temperature (C) or at the Hot Door Deployment temperature (H) when the door is closed during thermal balance is also provided in Table 3-36.
3.5.2.4.2 SoloHI Structural Thermal Model The SoloHI instrument module (SIM) DM is the Structural Thermal Model (STM) that shall be delivered to ESA for observatory-level vibration, thermal vacuum and separation shock tests. The SIM DM shall include a flight-like structure, brackets, instrument mounts, front baffle assembly, and baffle cover door assembly with mass simulators for the SIM telescope, the interior baffle assembly, and SCE. Flight-like thermal blankets, operational and survival heaters, and thermal sensors will also be integrated on the SIM DM for thermal balance tests at the observatory level. The SoloHI STM that is delivered to ESA will also include mass simulators for the SPS with heaters to simulate the average power dissipation and the SPS to SIM wire harness.
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Table 3-36. Environmental Test Matrix for SoloHI Hardware Levels of Assembly at Unit Level or Higher
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The SIM sine and random vibration test will verify that the SIM structure, front baffle assembly, brackets, instrument mounts, and the baffle cover door mechanism can survive the launch environment and that the SIM F1 front baffle alignment to the SIM mounting interface satisfies its ground alignment shift requirement. Since the SIM DM will undergo additional vibration testing at the observatory level, the vibration test will be conducted at protoflight levels and durations. The SIM DM thermal balance test will validate the SIM DM thermal model and demonstrate that all operational and survival heaters are operating nominally. The SIM DM thermal cycling test will demonstrate the survival of the SIM structure, mounts, and door mechanism, and will also demonstrate that the SIM F1 front baffle alignment to the SIM mounting interface satisfies its ground alignment shift requirement after exposure to the survival and operational cycles. The SIM DM baffle cover door will be deployed using the flight-like mechanism at the Cold Door Deployment plateau during the Thermal Balance in thermal vacuum.
3.5.2.4.3 Electronics The SPS and the SIM telescope with the camera electronics FMs will undergo thermal bakeouts and thermal vacuum cycling before integration on the SoloHI science payload. The SIM telescope module shall include the lens barrel assembly, the focal plane assembly with the detector package and DRB, and the SCE unit mounted on a thermal test fixture. The SIM telescope module DM will undergo thermal vacuum cycling tests to show that the camera electronics module will survive and satisfy all functional/performance requirements across its operational temperature range. The SCE DM will undergo a full EMI/EMC test program to verify that the EMI/EMC requirements are satisfied before the FM is built. The SCE DM and the camera electronics DM will also verify the functional and performance requirements at room temperature. GEVS and NASA Gold Rule requirements for a number of electronics failure-free hours of performance at on-orbit operational temperature and over a minimum number of temperature cycles are designed to identify electronics workmanship and infant mortality issues. The SoloHI flight electronics, including the SPS and the SIM telescope with the SCE and DRB, are required to operate for a minimum of 200 failure free hours in thermal vacuum in the operational temperature range before launch and to operate for a minimum of 100 burn-in hours in thermal vacuum at the hot operational temperature before launch. These flight electronics are also required to be operated over 12 operational cycles in thermal vacuum before launch. In addition, there is a requirement for flight electronics to demonstrate 1000 failure-free hours in air and in vacuum before launch. Table 3-37 compares the number of cycles and the best case number of failure-free hours in thermal vacuum for the baseline SoloHI environmental test program to the corresponding GEVS requirement. The required number of failure-free hours in air for the SoloHI flight electronics is also estimated to reach the 1000 failure-free hours in air and in vacuum before launch. The SoloHI flight electronics shall meet or exceed all of the above requirements. In fact, the SoloHI flight electronics will satisfy the failure-free hours in thermal vacuum and the burn-in hours requirements in the baseline SoloHI environmental test program in the box-level and payload-level thermal vacuum tests before the SoloHI science payload delivery for integration on the Solar Orbiter observatory.
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Table 3-37. Cycles, Failure-Free Hours and Burn-in Hours for SoloHI Electronics
3.5.2.5 Functional, Performance, and Characterization Test Program The SoloHI subsystem functional and performance requirements are defined in the SoloHI Instrument Specification (ISPC). The SoloHI functional, performance and characterization tests are shown in Table 3-38 for the SoloHI hardware levels of assembly from Subassembly level to the Observatory level. The table is divided between the SoloHI subsystem Flight Models and the Qualification/Risk Reduction Models. . Special performance tests, such as performance tests in thermal vacuum, alignment checks, straylight tests, and life cycle vacuum test, are distinguished from the standard performance test in air in these tables. Optical and straylight model validation will be performed with the end-to-end optical performance and characterization tests. Initially, the SIM telescope and baffle assemblies will be characterized at the subassembly level and continue to the end-to-end characterization at the science payload level. Optical performance tests by the vendors and at NRL will verify the throughput, imaging and straylight performance for the baffle and lens assemblies. The detector performance tests will verify its quantum efficiency, dynamic range, resolution, and noise. Additional end-to-end performance and characterization tests that will be performed under vacuum include vignetting, radiometric (responsivity), image quality, wavelength range, stray light and flat field test. End-to-end tests will use the flight instrument and camera electronics. All calibrations will be directly traceable to NIST using secondary standards. End-to-end testing is performed with the MOC/POC using the same ground S/W and databases to be used for flight. Rehearsals and training are conducted in the test flow. The baseline SoloHI Comprehensive Performance Test (CPT) is performed using instrument GSE after the initial science payload integration at NRL to characterize the baseline performance and verify all hardware and flight software functions, modes, and operations The CPT is repeated at the end of the science payload environmental tests at NRL. At the observatory level, the CPT will be repeated before integration on the Solar Orbiter spacecraft at Astrium. All operations are conducted in a clean room, or localized clean tent at the observatory level, that satisfies the facility class requirements (per FED-STD-209E) specified in the SoloHI Contamination Control Plan.
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Table 3-38. SoloHI Verification and Characterization Test Matrix
Variance tests shall be performed periodically throughout the SoloHI instrument and Solar Orbiter observatory integration and test program, and are the primary type of performance verification that is performed after the SoloHI instrument is delivered to the observatory for integration. Variance tests measure the change in the instrument state or performance from the initial baseline measurement, and determine if the measurement change from transportation and handling, testing or exposure during the environmental test program, satisfies the related
3-85 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 performance requirement. Variance tests are performed for alignment, optical throughput, spatial resolution, and surface cleanliness (both particulate and non-volatile residue). In addition, flight software tests are repeated throughout the instrument and observatory I&T program to show that the performance of the electronics, camera, and thermal system has not changed after the baseline flight software tests on the SoloHI science payload. The flight software variance tests include abbreviated functional tests, full functional tests, and camera performance tests. Table 3-39 shows the variance tests for the SoloHI science payload during the instrument and observatory I&T. The baseline measurement (marked as “B” in Table 3-39) is taken for each of these variance test metrics for the SoloHI science payload, the SIM and the SCE before the start of the Instrument environmental testing. Variance tests continue to be performed during the instrument and observatory I&T program. Table 3-39. Variance Test Matrix for SoloHI Flight Hardware
3.5.2.6 Contamination Control and Cleanliness The SoloHI contamination control requirements and practices are adopted from SECCHI and LASCO, in which no serious optical performance degradation has been observed on-orbit. The SoloHI Contamination Control Engineer has the prime responsibility for the cleanliness of the SoloHI science payload from procurement/fabrication through assembly, integration and test to launch. The SoloHI Contamination Control Plan (CCP) shall define the SoloHI surface cleanliness levels (particulate and molecular) for internal and exterior surfaces at the baseline mission EOL. The SoloHI CCP will derive the SoloHI surface cleanliness budget throughout the SoloHI AIT flow including subsystem delivery for science payload integration, instrument delivery for spacecraft integration, launch and the end of the baseline science mission. The cleanliness inspection level requirements will be subsequently defined for each of these AIT periods. The outgassing rate requirements for each SoloHI subsystem are derived from the molecular surface cleanliness requirements using a three-dimensional molecular transport model. Materials and process list shall verify that material selection requirements for total mass loss (TML) of <1.0%, collected volatile condensable material (CVCM) of <0.1%, and other material restrictions are satisfied. The SoloHI contamination control program will adopt design strategies for composites use, venting, purging, cleaning during I&T, and accessibility for inspection and cleaning that have been successfully applied to LASCO and SECCHI. The CCP will describe the facilities class and
3-86 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 air molecular cleanliness, garments, personnel training, bagging and monitoring to maintain the required surface cleanliness during I&T. Special precautions are addressed for environmental testing and transportation. All flight hardware will undergo thermal bakeouts with TQCM and Residual Gas Analyzer (RGA) monitoring. The SoloHI contamination control plan will achieve overall instrument cleanliness by component cleaning and bake-out, optical instrument assembly/ storage/test in a Class 100 cleanroom, a continuous GN2 purge with closed door until just before launch, and a Class 10,000 white room with a localized clean tent for S/C I&T activities. 3.5.2.7 Facilities and Ground Support Equipment The SoloHI instrument performance will be tested and characterized in the dedicated NRL coronagraph test facilities (Korendyke et al. 1991). The laboratory contains an 11m beamline optical test chamber and Class 100 cleanroom and is ideally suited for this application. -14 Additional baffling will be added to the chamber to allow the testing of stray light to ~10 B/Bs. Similar modifications were implemented for the successful SECCHI/HI instrument end-to-end stray light test. This test was the first test to successfully achieve this level of sensitivity. The chamber is equipped with collimating optics, a precision instrument pointing table and necessary light sources. The laboratory is equipped with optical benches, theodolites, alignment telescopes, optical flats, and light sources. Component transmission and reflectivity will be characterized using a Cary spectrophotometer and spectroradiometer.
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4. MANAGEMENT 4.1 Management Processes The Principal Investigator (PI), Dr. Russell A. Howard of the U.S. Naval Research Laboratory, leads the SoloHI investigation. Figure 4-1 shows the SoloHI organization. Table 4-1 and Table 4-2 show the roles and responsibilities of the institutions and team members. Our team has a proven success record from the recent successful delivery, launch, and operation of the STEREO/SECCHI Instrument suite and is committed to achieving the following objectives for the SoloHI Investigation: Conducting the scientific investigation described in Sections 2 and 3 of this report. Designing, developing, fabricating, testing, calibrating, integrating, and operating the instrument to acquire the necessary observational data. Managing the personnel, resources, and interfaces to accomplish the project on schedule, within budget, and in a manner that minimizes risk and maximizes the science return on expenditures. Meeting NASA‟s E/PO and small/small disadvantaged business contracting goals. Performing the mission operations and data analysis activities after launch throughout the MO&DA phase. Our project management process has evolved from a successful series of programs at NRL. Years of experience in flight hardware, software, and ground data systems, combined with our understanding of the GSFC approach to space missions, enable us to accomplish this investigation at modest cost and minimal risk. The PI is committed to a successful mission within the allowed budget and is fully prepared to recommend termination to NASA if this cannot be achieved. We maintain clear lines of authority, responsibility, and reporting (Section 4.1.7). The Project Manager (PM) (Figure 4-1) reports directly to the PI to ensure integrated management and decision-making processes. The Project Management Office (PMO) includes the PI, the PM, Dr. Simon Plunkett (NRL), the Systems Engineer (SE), Dr. Timothy Carter (Praxis, Inc.), and the Mission Assurance Engineer (MAE), Mr. Donald McMullin (Space Systems Research Corporation). Our plan is built on the dedication and personal commitment of each team member, with the full support of his/her institution. Each team member brings direct experience with relevant science objectives and focused spaceflight instrumentation. A Work Breakdown Structure (WBS) dictionary (Appendix A) per NPR 7120.5D, Appendix G, defines the limits of authority relative to requirements, cost, and schedule. It serves as the main formal management tool to control the project and support Earned Value Management (EVM) implementation (Section 4.1.7.2). The PM uses a proven Project Management and Control System (PMCS) evolved from NRL‟s STEREO/SECCHI, Hinode/EIS, and SOHO/LASCO/EIT projects. It provides direct accountability for essential project elements to achieve the project objectives on time and to specified cost, quality, and performance metrics (Section 4.1.7.1).
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We have budgeted and baselined adequate technical and programmatic reserves. Their allocation is centrally managed (Section 4.1.3) and formally distributed under a Configuration Management Plan (CMP). The PMO briefs senior NRL management monthly during the project review process. These meetings facilitate the institution‟s role of technical support and providing an “early warning” system in detecting and resolving emerging problems. NRL leads SoloHI project management, and brings extensive experience in successfully managing multi-institutional projects as part of multi-national missions. This includes a 60- year history with instrumentation and spaceflight projects on a scale equal to or larger than SoloHI (e.g., SECCHI, EIS, LASCO).
Figure 4-1. Our Streamlined Organizational Structure Provides Clear and Simple Lines of Project Accountability and Communication
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Table 4-1. The PI has Implemented Well-Defined Roles, Responsibilities, and Commitments With an Experienced Institution, the NRL, to Define, Implement, and Execute the SoloHI Project
Mission Institution Relevant Experience Unique Capabilities Responsibilities Naval Research Laboratory, Washington, DC . Numerous sounding rocket . Responsible institution for . Leads SoloHI and spaceflight the SoloHI instrument, and instrument design, missions the home institution for the development, and PI, PM, and SE. test . Complete H/W design, fab, assy, . Lead for project . Leads SoloHI . SOHO/LASCO and test capability management and reporting, instrument IAT with . SOHO/EIT mission assurance, the S/C . Class 10k contamination control, . HINODE/EIS cleanrooms, . Leads ISPC/ISDC system engineering, unique Class 100 development & . STEREO/SECCHI SSR&MA, and project-level TVAC and mission ops CM FUV/EUV . Leads MO&DA to . Design/Fab/I&T/Cal of calibration deliver scientific SoloHI and all electronics chamber with full data products H/W, S/W, optics, and environmental test structure, including facilities harnessing and EGSE Scientific Co-Is, relevant experience, and background are shown in Others Table 4-2
Table 4-2. SoloHI Key Team Members
Individual Relevant Experience and Responsibilities
SoloHI Project Management Office Key Personnel . Experience: PI STEREO/SECCHI and SOHO/LASCO . Responsibilities: Scientific success and implementation of SoloHI; leader of the science team and ultimate decision maker on cost reserve allocation; Dr. R. A. Howard, development of instrument performance requirements within allocated PI, NRL resources; delivery of measurements to NSSDC and coordination of activities to NASA SMD; definition of data products and active participation on the SoloHI SWG. . Experience: PM for STEREO/SECCHI Operations, Operations Scientist for SOHO/LASCO, SECCHI/COR2 Instrument Scientist Dr. S. P. Plunkett, . Responsibilities: Day-to-day management of the SoloHI project; CRM process; PM, NRL decision and requirements management processes; budget allocations and tracking; scheduling, team member subcontract establishment and management; contract compliance with NASA SMD. . Experience: SE STEREO/SECCHI . Responsibilities: P/L technical resource allocation and tracking; P/L interface Dr. M. T. Carter, definition and control; derivation of P/L environmental design and test SE, Praxis, Inc. requirements; leadership of trade studies; coordination of P/L requirements, interfaces, and operations to the spacecraft team; verification and validation processes; providing assistance to the I&T manager for test requirement.
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Individual Relevant Experience and Responsibilities . Experience: DPM STEREO/SECCHI, SS&MA lead for HREP mission on ISS, PM and SE for SOHO/SEM . PE Responsibilities: Ensure compliance of procured hardware with project requirements; ensure correct timing, resourcing, and sequencing of work effort; Mr. D. R. McMullin, communicate project information to team members and stakeholders; manage PE/MAE, SSRC and resolve issues or escalate to PI/PM; provide PM with go-no go decision points at key milestones. . MAE Responsibilities: Lead SoloHI safety and assurance activities; manage development of SS&MA program as per PAIP; chair MRB as needed.
JPL Key Personnel . Experience: Scientific analysis of solar and heliospheric data and comparison Dr. P. C. Liewer, with models Co-I, JPL . Responsibilities: Lead JPL Co-I; visualization and analysis of SoloHI images.
Other Key Personnel . Experience: Deputy PI of LASCO, Program scientist of SECCHI. COR2 Dr. A. Vourlidas, instrument scientist NRL . Responsibilities: Deputy PI; Monitor compliance to science requirements and support the development of program-level requirements. . Experience: PI of many rocket flights of HRTS and VAULT and PI of a new sounding rocket experiment, VERIS. Instrument scientist for HINODE/EIS. Dr. C. Korendyke, Consultant on SECCHI. NRL . Responsibilities: Instrument Manager; Ensures that SoloHI will meet the requirements. . Experience: Instrument Scientist for RAIDs and the Middle High Resolution Dr. J. Morrill, Spectrograph Investigation (MAHRSI) NRL . Responsibilities: Deputy Instrument Manager; Data Archive Support. . Experience: PM STEREO/SECCHI, ISS Configuration Manager Ms. R. A. Baugh, . Responsibilities: Technical support for budget allocations and tracking, Project Support schedule development and implementation, CM implementation. 4.1.1 Management Organization Figure 4-1 shows our organizational and institutional relationships. NRL, as the PI institution and recipient of the prime contract from NASA, has overall responsibility for the management of the investigation. There are clear reporting paths at all levels within the team. In accordance with the Section 10.5.1 of the AO, the overall SoloHI Investigation management, implementation, and execution is the responsibility of the NASA Goddard Space Flight Center (GSFC) Solar Orbiter PM. This includes responsibilities for SoloHI financial, procurement, technical management, and performance aspects within the overall requirements of the Solar Orbiter mission. All formal communications and agreements concerning Solar Orbiter technical and programmatic aspects are made between the ESA PM and the Solar Orbiter PM. All scientific aspects of the SoloHI investigation are coordinated by the NASA Solar Orbiter Project Scientist who is the formal interface for SoloHI scientific matters in liaison with ESA.
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4.1.2 Decision-Making Process The PI is directly responsible to NASA for mission success (ref. GPR 7120.3B). He has delegated a significant level of responsibility to the PM to define and execute integrated management processes. The PI is the final authority for changes affecting project scope while the PM is the final authority on allocating resources, schedules, and requirements among the Level-1 WBS elements. The PI recognizes that scientific, technical, schedule, or cost issues can arise at any time and has established a structured decision-making framework (Figure 4-2). 4.1.3 Management of Reserves and Margins Reserves are “rolled-up” to the project level with a >30% reserve maintained on mission phases B-D and 10% in phase E. Reserves fall into two major categories, technical and programmatic, and they are allocated to (i) fit the funding profile, (ii) reflect the design maturity level by phase, and (iii) match the risk level associated with each project phase. The PMCS (Section 4.1.7.1) tracks cost and schedule variances against the planned budget. These data, when coupled with earned value metrics, characterize work progress towards established milestones. Managing the variances lies with each subsystem, yet if cost or schedule variances deviate from the budget baseline, the CM methodology ensures that overall technical and programmatic reserves are managed. Our reporting and corrective action approach ensures that reserve is used sparingly, as part of the overall risk mitigation plan discussed in Section 4.3, and not simply as a stopgap measure. Technical Reserves. The SE manages technical reserves. Initially each subsystem is allocated resources based on a combination of current best estimate plus uncertainty. Uncertainty is managed initially as the subsystem reserve and is available to use when needed. The SE holds margins above reserves. Margins are allocated only through a formal process whereby the subsystem lead justifies the need for the additional resources, the steps taken to prevent the need, and the consequences if the resources are not granted. Using these data, the SE decides on the additional allocation of resources. The SE determines the rate of margin usage and the point in the schedule where he feels comfortable releasing reserves. The SE reports resource usage on a monthly basis in the form of time history plots. Programmatic Reserves. The PM manages programmatic reserves. Funded schedule reserves are built into the instrument development schedule. The PM makes his decision based on the justification provided by the SE and the consequences to the instrument schedule and delivery of the payload to the spacecraft vendor. The PI makes the final decision to release cost reserves based on input from the PM and the SE and the prior rate of usage of reserves from all team members. Using the process described in Section 4.1.2, the PM reviews any request for additional funding and makes a recommendation to the PI. A maximum rate of usage for cost reserves is established based on the phase of the mission, the previously used reserves, and the total developmental costs to complete. Any usage of reserves above this rate is grounds for a descope action.
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Figure 4-2. A Systematic Decision-Making Process Supports the Use of Margins and Reserves
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4.1.4 Communications Process Our management approach relies on continuous and open communications. Although decisions are made in a structured manner, ideas are shared openly. Decisions are documented in meeting minutes and technical memoranda and then incorporated into the appropriate documents. Project members are linked via e-mail and a secure project website where all specifications and project documentation are maintained. The website includes database tools to implement a complete system level verification process. Action items are tracked via a web-based Action Item Management System (NPR 7120.5D). 4.1.5 Staffing Plan 4.1.5.1 Principal Investigator Program authority is delegated from the Associate Administrator for the Science Mission Directorate (AA/SMD) through the GSFC Center Director to the Solar Orbiter Project Manager within the Flight Projects Directorate at GSFC to the instrument PI. The PI is responsible for the overall success of the SoloHI Instrument and is accountable to the AA/SMD for scientific success and to the GSFC Center Director for programmatic success. He is NASA‟s primary point of contact to lead and direct the project‟s scientific, technical, and business efforts and is fully accountable to NASA for all decisions made and actions taken, including accomplishment of the mission within defined cost and schedule constraints (ref. GPR 7120.3B). This includes approving all instrument specifications, advising the PM in development and fabrication, participating in final calibration, developing the operations plan, and analyzing and interpreting data. He is committed to providing the maximum science with low cost, schedule, and minimal technical risk, and achieves these objectives using the PI mode of management with clear lines of authority, in-place program controls, defined roles and responsibilities, and NASA insight and involvement. He makes the final decision on personnel selection, has the authority to make changes in the technical or scientific baseline within NASA guidelines, and is responsible to implement decisions and recommendations of NASA reviewers. He must approve any contract scope change or disbursement of project reserves. 4.1.5.2 Project Manager The PM is delegated complete authority by the PI to accomplish the defined performance, schedule, and cost goals. The PM maintains technical, programmatic, and fiscal interactions with the PI and NASA. He negotiates commitments with supporting and partnering organizations and institutions. He establishes and implements project policies, and prepares the Project Plan. He ensures mission elements are developed into a consistent set of requirements that support the Level 1 baseline science objectives agreed upon by the PI and science team. The PM ensures that clear lines of communication are defined and used across the SoloHI team. 4.1.6 Systems Engineering Management Plan The SE establishes a systematic and disciplined system engineering process during Phase A. It addresses functional, physical, and operational performance requirements during the instrument‟s planned life cycle and follows the guidelines of NPR 7123.1A. The process is compliant with the GSFC Gold Rules and our test program is based on GSFC-STD-7000. Our approach includes formal requirements development, baseline management, verification compliance, technical performance metrics, peer reviews, detailed schedules, and weekly status meetings. The process includes applicable design criteria, analysis, support, test procedures,
4-7 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 lessons learned, environmental and safety procedures, and a collaborative engineering methodology. During Phase B, the SE performs system decomposition and design until the “design-to” specifications of all lower-level configuration items are released. The SE audits the H/W and S/W designs for compliance to all higher level baselines. The SE ensures and documents requirements traceability. The SE is accountable for system trades, requirements decomposition, and developing specifications, design verification plans, and ICDs. All requirements are formalized, documented, traced, and verified using a requirements traceability tool. In a similar manner, the system engineering activities continue during Phase C/D. 4.1.7 Management Approach Our project management processes incorporate NASA NPR 7120.5D guidelines, along with successful practices and “lessons-learned” on STEREO/SECCHI, SOHO/LASCO, and Hinode/EIS. Our senior team members have experience from recent successful spaceflight missions (SECCHI, LASCO, EIS). We use system engineering processes per NPR 7123.1A that emphasize a clear, but limited, set of science goals and objectives synthesized via a defined process into performance requirements and specifications (ref. NASA SP-610S). We deliver a Systems Engineering Management Plan (SEMP) per NPR 7123.1 at the SRR. It describes the overall approach for system engineering from early design through product realization. It also describes how performance verification and the technical management process are accomplished. Section 4.1.5 describes the roles and responsibilities of the PI and the PM. The PM delegates authority to the SE to execute the system engineering process. The SE works with the science team to ensure science goals are synthesized into project performance specifications. 4.1.7.1 Project Management Control System During Phase B, the PM establishes an integrated PMCS to monitor and assure compliance with cost and schedule baselines. Our PMCS provides the following: (i) integrated cost and schedule baseline; (ii) work authorization; (iii) management reports displaying cost, funding, and schedule status vs. baseline plans; (iv) actual and forecasted cost and schedule status against the Performance Measurement Baseline (PMB); (v) a clearly documented audit trail of changes to the PMB through the WBS; (vi) identification of problem areas requiring management actions; and (vii) identification of cost issues needing corrective action. 4.1.7.2 Earned Value Management Our PMCS approach integrates EVM as a project management tool versus a financial management tool. NRL‟s heritage EVM approaches (Upper Stage, WindSat) incorporate ANSI/EIA-748A processes. Major EVMS activities include: (i) planning the entire scope of work to completion; (ii) integrating program scope of work, schedule, and cost objectives into a baseline plan against which accomplishments are measured; (iii) assessing accomplishments objectively at the work performance level; (iv) analyzing significant variances from the plan and forecasting impacts; and (v) providing data for decision-making and management actions. We develop a CADRe EZ during Phase A; subsequent submissions follow NPR 7120.5D, Appendix E. 4.1.7.3 Requirements Process The PI and SE define the Level 1 requirements and flow downs to Level 4 during Phase A in the SoloHI Science Requirements Document (SSD-RQT-SOLOHI-0001). It shows how each requirement is measured and verified and is used to confirm that all systems meet all
4-8 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 performance requirements. Mission requirements test and verification are initiated during Phase A, baselined at SRR, and maintained throughout the project. The mission design process follows NPG 7120.5D guidelines, including mission analysis and definition; requirements analysis and allocation; design, development, fabrication, and manufacturing; test and verification; and operations. Review milestones allow external readiness assessments by NASA. 4.1.7.4 Mission Assurance Process The MAE establishes an organized SSR&MA program and documents it in a Product Assurance Implementation Plan (PAIP). It includes our approach to meeting mission assurance requirements of the Solar Orbiter Project Instrument MAR, and it addresses the CMP and CCB. The final PAIP is submitted at the Confirmation Review. NRL manages the SSR&MA program and reviews team member processes. We comply with NPR 8705.4 for safety requirements, and a systems safety engineer is involved with all aspects of the H/W, S/W, and GSE, reviewing them for safety issues/concerns. We implement a tailored approach to meet NPD 8720.1 reliability program requirements for Class C missions and it includes reliability engineering processes (ref. NPR 7120.5D). The MAE is responsible for all project-level SSR&MA plans. NRL and supporting team members are responsible for their local processes, and our SSR&MA approach is subject to NASA review. We use existing processes and procedures to address SSR&MA elements including contamination control and cleanliness. We maintain a closed-loop system to identify and report on non-conformances, and ensure that corrective action is taken. All non-conformances are dispositioned via a Material Review Board. Failure reporting starts early in the life cycle. A Failure Review Board dispositions all failures. 4.1.7.5 Acquisition Process Procurement responsibility resides with the PM. Make/buy decisions are based on trade studies led by the SE. Critical subcontracts are subject to PM approval. Each acquisition is managed by a contracting official. Procurements are selected via competitive solicitations meeting the Federal Acquisition Regulation (FAR) and NASA FAR Supplement. Suppliers must have a proven record of meeting cost and schedule constraints. In-place support contracts provide many of the technical services, suppliers, and materials. Conformance is monitored via design reviews, audits, and acceptance tests. QA personnel visit suppliers to verify that levied requirements are being met. Deviations in schedule, cost, or performance are flagged for management attention. The PM reports on major procurements and contract status monthly to the PI. 4.2 Schedule Table 4-3 provides the key milestones for the SoloHI program. The detailed SoloHI Integrated Master Schedule is provided in a separate Project and PDF files. The Schedule Management Plan for the Solar Orbiter Heliospheric Imager (SoloHI) (NRL, SSD-PLN-SOLOHI-0004, delivered separately) defines the process that will be implemented for the management of the SoloHI Instrument Integrated Master Schedule (IMS).
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Table 4-3. SoloHI Key Milestones
Milestone Date
Kick-Off Meeting 04/24/2009
Instrument Design Status Review (IDSR) 10/08/2009
Draft Concept Study Review (CSR) Delivery 08/16/2010
Instrument Systems Requirement Review/System Design Review 09/14/2010 (SRR/SDR)
Draft 2 CSR Delivery 12/31/2010
Instrument Technical Status Review 10/25/2011
Final CSR Delivery 12/31/2011
Instrument Preliminary Design Review (PDR) 08/15/2012
Instrument Critical Design Review (CDR) 06/17/2013
Instrument Structural Thermal Model (ISTM) Delivery 07/15/2013
Instrument Electrical Engineering Model (IEM) Delivery 01/15/2014
Instrument Pre-Environmental Review (PER) 05/29/2014
Instrument Pre Ship Review (PSR) 10/02/2014
Flight Instrument Delivery 1/30/15
4.2.1 Critical Path The current critical path is identified in the schedule as the SoloHI Camera Electronics (SCE) development. This is primarily due to the extended time allotted for EEE parts procurement (13 months). This duration is a worst case estimate and is included to account for any possible delays due to failures during part testing, GIDEP alert issues, and supplier issues. The SoloHI secondary critical path is the Focal Plane Assembly (FPA) development. The FPA has a full up qualification test program in line prior to its flight unit fabrication. The qualification model (QM) FPA cannot be completed until the engineering model (EM) APS detectors are delivered. Any delays in the APS detectors would delay the FPA qualification program and push out the flight program. Because of this, options are in place to develop the flight FPA in parallel with the QM FPA. Because the FPA is based on a heritage design, the risk in fabricating these in parallel is minimal, with the radiator size being the biggest concern. If the qualification test results identify a need to modify the radiator size then a new radiator can be fabricated in parallel with the SoloHI integration and swapped out prior to instrument level environmental testing.
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4.3 Risk Management 4.3.1 Top Risks and Mitigation We use legacy Risk-Informed Decision Making (RIDM) approaches that have been developed and used on previous projects (STEREO/SECCHI, Hinode/EIS) to establish a Continuous Risk Management (CRM) process compatible with NPR 8000.4A and NPR 7120.5D (NID), including CRM-complementing processes (e.g., Risk-Based Acquisition Management and Risk-Based Decision Support). The CRM process is a continuous and iterative five-step process that (i) identifies; (ii) analyzes; (iii) plans; (iv) tracks; and (v) controls risk throughout the project life cycle. It includes documentation and communication, with pre-defined triggers and off-ramps. A Risk Management Plan (RMP) will be submitted at PDR. 4.3.2 Risk Management Approach RM started during Phase A, and will continue throughout the project life cycle. The RM process includes risk identification, acceptance, action planning, prioritization, and modification/closure. Risk identification begins with a statement of the form „IF {event}, THEN {risk to schedule, budget, or technical/science}‟. All team members contribute to identifying risks and submitting them to the RM Board (RMB) in monthly reports. The RMB meets monthly and includes the PI, PM, Deputy PI, Systems Engineer, and Instrument Manager. Other subsystem leads are invited to attend RMB meetings for risk acceptance, prioritization, and modification/closure in their area of responsibility, but they are not required for RMB approval. Risk acceptance occurs if a majority of the RMB agrees to include a candidate risk on the project risk list. The RMB formally accepts each risk and assigns values for risk likelihood (Lf), risk impact (Cf) if the risk occurs, timeframe, and the RMB action (mitigate, watch, accept, or research). Risk action planning is performed by a RMB member who is designated at the time of risk acceptance to generate a mitigation, watch, or research plan with schedule deadlines before the next RMB meeting. Risk mitigation plans include contingency plans with trigger dates and/or events, if the primary risk mitigation plan is not working by those dates. Risk action plans are approved by the RMB. Risks are prioritized by the risk exposure value (defined as the product of Lf × (Cf-0.5)), then sorted by their risk impact. Prioritized risks are reviewed in Top 10 lists or in Risk Tables. Risk modification/closure occurs at the monthly RMB meetings. The RMB updates risk exposure values based on new information and closes risks if the risk likelihood falls below a threshold value. Risks can be accepted if the risk does not affect a Level-1 science requirement and if the risk exposure cannot be reduced further. 4.3.3 Risk Management Tools The SoloHI RMS uses a prioritized risk list and the risk table as decision aids to allocate project resources to reduce the project‟s risk exposure. The risk list documents, prioritizes, and manages the programmatic and technical risks with their associated action plans in an Excel spreadsheet. The risk table uses a 5 × 5 Likelihood – Impact diagram to map the risk exposure. The risks in this table are categorized as „High‟ (red), „Moderate‟ (yellow), and „Low‟ (green). The RMP provides a tailored budget, schedule, and science/technical definition for each of the five impact levels. Risks are reported to NASA on a monthly basis. 4.3.4 Top Risks Table 4-4 shows the current risks (current as of December 2011) ranked in order of descending risk exposure, then sorted by the highest risk impact if two risks have the same risk exposure
4-11 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 value, with the action identified by the RMB shown for each risk. Figure 4-3 shows the corresponding current risk matrix. This prioritized risk list and matrix are updated on a monthly basis. Table 4-4. SoloHI Risk Table Based on Risk Exposure
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Figure 4-3. SoloHI Risk Matrix 4.3.5 Managing New Technology Risks The SoloHI Technology Development Plan (NRL, SSD-PLN-SoloHI-0001) has identified the SoloHI APS detector as the only SoloHI subsystem or part at a NASA technology readiness level (TRL) of ≤6 (system/subsystem model or prototype demonstration in a relevant environment). The current TRL for the SoloHI APS detector is equal to 5 (component or breadboard validation in a relevant environment). As part of the Phase A effort, performance requirements were developed and a trade study of device candidates was completed. See Appendix E for details on the detector trade study and backup plans. The SoloHI APS detector is based on the Sarnoff Minimal array detector family, including the 3T and 5T pinned photodiode (PPD) pixel designs. The Sarnoff Minimal array detector performance has been cited in numerous papers over 5 years of government-funded and scientific-funded research. (See the SoloHI Technology Development Plan for these papers.) The Sarnoff Minimal array detector performance has been tested at operating temperatures from room temperature to -100°C and has continued to perform well after thermal vacuum cycling testing and radiation exposure. The flight design for the SoloHI APS detector is based on a large format (4k x 4k) front- illuminated detector with a 10 µm 6T PPD pixel design and a high readout rate of 4 Mbps in parallel for the top and bottom detector halves. The 6T PPD pixel in the SoloHI APS detector has added an extra transistor to the heritage Sarnoff Minimal 5T PPD pixel to increase the sense node capacitance and thereby increase the linear full well. The SoloHI APS detector is required to survive and meet all functional and performance requirements after exposure to the radiation and thermal environments. The SoloHI detector thermal requirements include -85°C to -50°C for the operational case and -110°C to 80°C for the survival case. The detector must be
4-13 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 demonstrated to survive a radiation environment with a radiation dose margin (RDM) of 2x above the mission radiation environment. This would result in a test requirement of 96.8 krad for total ionizing dose (TID) and 2.38e11/cm2 for non-ionizing energy loss (NIEL) based on 10 MeV protons. The SoloHI Technology Development Plan (NRL, SSD-PLN-SoloHI-0001) describes the plan to raise the TRL from 5 to 6 (system/subsystem model or prototype demonstration in a relevant environment) by the SoloHI instrument PDR. The SoloHI APS detector TRL can be promoted by demonstrating that: The detector design meets all functional and beginning of life (BOL) performance requirements. The detector design can survive and meet all functional and end of life (EOL) performance requirements with the following exceptions: 99.7% of pixels shall meet performance after exposure, and the dark current is allowed to increase by 12x after exposure. The performance of the SoloHI APS detector will be tested across the operational temperature range of -85°C to -55°C. The Sarnoff Minimal family has already been well documented to survive and still perform nominally after exposure outside of the SoloHI survival temperature range of -110°C to 80°C. Development models for the SoloHI APS detector will be designed, fabricated, and tested for the SoloHI detector technology development program. The Sarnoff Sandbox VI detector, which includes a front-illuminated, 410 x 410 pixel format with the SoloHI 6T PPD flight pixel design but does not include the flight high speed readout circuitry, has been designed and fabrication started during Phase A. The Sandbox VI detector has been delivered to NRL in 2011 and pre- radiation tests at operational temperatures, radiation tests, and post-radiation tests at operational temperatures will be performed during Spring 2012. NRL plans to hold a SoloHI detector technology development review in late June 2012 to show that the detector TRL can be formally promoted. If the Sandbox VI does not satisfy the required functional and performance requirements, a SoloHI FRB will determine how to proceed. If the requirement violation is small, the SoloHI error budget from the top-level photometric accuracy requirement can be revised to show that the reduced detector performance is allowed if other requirements in the error budget are tightened. The revised SoloHI error budget must show positive margin for all derived instrument and detector requirements. If the detector requirement violation is large such that a change in the SoloHI error budget will not satisfy the photometric accuracy requirements, the SoloHI FRB will trigger the backup plan for procuring the detector from a vendor other than Sarnoff. The SoloHI detector backup plan was described in detail in the SoloHI SRR request for action (RFA) response to RFA MLB-002. All SoloHI backup detector candidates are off-the-shelf detector procurements that are either already at a TRL of ≥6 or whose BOL performance satisfies the SoloHI requirements, but require a delta radiation test to achieve a TRL of 6. The SoloHI backup plan describes how the backup detector will be at a TRL of 6 by the SoloHI PDR. 4.4 Government Furnished Property, Services, and Facilities Per the FOSO AO NNH07ZDA003O, we expect that the following services will be provided by NASA:
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ESA Interface: The NASA Solar Orbiter Project Office at GSFC will handle all interfaces with ESA. There are no other anticipated government furnished property or facilities required. Based on previous experience, the following services may be needed from NASA/GSFC: Consultation on material selection including low outgassing materials and lubricants. Thermal consultations on absorptance and emittance measurements of thermal coatings. Application of conductive thermal paint on selected parts. Radiation and part selection consultation for this unique environment. Engineering advice on unusual problems that may arise during the development. 4.5 Furnished Property, Services, and Facilities The PI is responsible for mission design, development, test, flight operations, and coordinating the work of the team members, contributors, and Co-Is. For the scientific investigation and the data verification tasks, the PI uses the Co-Is identified in Table 2-14. 4.6 Reporting and Review Plan Our heritage reporting processes support project management and provide consistent feedback to NASA. The PMCS maintains databases of the cost, schedule, and PMB, and NASA receives a Monthly Project Management Report. The PM prepares and submits monthly (533M or equivalent) financial reports pursuant to NPR 9501.2B, using the WBS (Appendix A). Financial management reporting is provided at WBS Level 2. Reports are provided for first-tier contracts using NASA guidelines. The PM prepares and the PI provides contract funding profiles and explains variances between projected and actual costs. Monthly status reporting to the NASA Solar Orbiter Collaboration project office will also be maintained. In addition to cost, the PI and PM will provide the current status of each element of the project, schedule, EVM reports, and action items. The PM establishes a rigorous technical review process for all aspects of the project. For all reviews, the SE presents designs, test plans, and verification plans. The review team develops specific recommendations, action items, and concerns that are tracked, and closures are presented at subsequent reviews. NASA personnel are invited to all peer reviews. Peer reviews are conducted consistent with the SoloHI Instrument Review Plan (SSD-PLN-SOLOHI-0005). These are convened by the SE as part of the process leading up to the formal reviews and outside technical experts engage in round table reviews at key development stages. Notes and action items are taken and are presented at the formal project reviews.
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Appendix A. WORK BREAKDOWN STRUCTURE AND DICTIONARY Table A-1. Work Breakdown Structure and Dictionary
Responsible WBS WBS Task Name and Description Manager / No. Organization Project Management: This summary element includes business and administrative planning, organizing, directing, coordinating, analyzing, controlling, and approval processes used to accomplish overall project objectives, which are not associated with specific hardware or software 1.0 elements. This element includes project reviews and documentation, non- PM/NRL project owned facilities, and project reserves. It excludes costs associated with technical planning and management and costs associated with delivering specific engineering, hardware and software products. All activities must be in compliance with NPR 7120.5D. Project and Mission Management: Includes program/project plan, IT security plan, and sub-plans of Part 3, NPR 7120.5D that are not addressed in WBS 2.0. Includes program/project planning; interface with NASA management; 1.1 financial control; production of cost estimates; management of project PE/NRL reserves; project reviews; project documentation; CDRLs; coordinate with the SE to develop continuous risk management processes and plans; and ITAR management. Business, Contract, Resource, and Procurement Management: Lead and manage the business and resource control processes for the Project. Includes 1.2 Project resource planning and control activities; contract initiation and PE/NRL management; procurement management; and provide configuration management support to the project team. Schedule/Planning Management: Lead the development and maintenance of 1.3 the Project schedules and the Project performance measurement and PE/NRL reporting system (SoloHI tailored Earned Value Management system). Systems Engineering/Mission Analysis: This summary element includes the technical and management efforts of directing and controlling an integrated engineering effort for the project. This element includes conducting trade studies, the integrated planning and control of the technical program efforts of design engineering, software engineering, specialty engineering, system architecture development and integrated test planning, system requirements 2.0 SE/NRL writing, configuration control of technical elements, technical oversight, control and monitoring of the technical program, and risk management activities. Documentation products include requirements documents, interface control documents (ICDs), Risk Management Plan, and master verification and validation (V&V) plan. Excludes any detailed design engineering costs associated with delivering the specific Payload H/W and S/W products.
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Responsible WBS WBS Task Name and Description Manager / No. Organization Systems Engineering Management: The management efforts associated with directing and controlling an integrated engineering effort for the project. Includes the efforts of the Systems Engineer (SE) who leads the Project’s overall system architecture, definition and engineering functions. This element includes the efforts to define the instrument payload and the science operations center ground data system, conducting trade studies; the integrated 2.1 planning and control of the technical program efforts of design engineering, SE/NRL specialty engineering, human rating, system architecture development, and integrated test planning, system requirements writing, configuration control, technical oversight, control and monitoring of the technical program, and risk management activities. Also includes the development and execution of an integrated technical program, as well as the system requirements document (SRD); and ICDs. Structural Analysis: Provide structural engineering support to the Payload 2.2 System development. This includes the development and periodic updating of SE/NRL all required system level structural analyses. Thermal Analysis: Provide thermal engineering support to the Payload 2.3 Systems development. This includes the development and periodic updating of SE/NRL the instrument system level thermal model. Contamination Control/Materials Engineering: Provide contamination and materials engineering support to the Payload System development. Develop 2.4 SE/NRL contamination control plan and review contractor contamination control procedures. Support integration and test activities at all levels of development. Project-Level Reliability, Verification and Validation: Develop validation strategy, create and review requirement linkages, develop verification strategy, help with project level scheduling and resources planning, monitor progress of 2.5 test program, report status to management, document verification and SE/NRL validation plans, track anomalies, compile and analyze trend data, and organize and attend verification and validation peer reviews. Includes the System Validation (V&V) Plan. Project Software Engineering: Lead the system software architecture effort, and function as the Project S/W Systems Engineer. Includes, development of 2.6 SW policies and practices; defining software requirements; SW design; SW SWE/NRL implementation; SW test issues; flight/ground trade-offs. Documentation products include: Software Management Plan. Systems Safety, Reliability, and Mission Assurance (SSR&MA): The technical and management efforts of directing and controlling the top-level safety and mission assurance elements of the project. This element includes design, development, review, and verification of practices and procedures and project 3.0 MAE/NRL success criteria intended to assure that the delivered payload meets performance requirements and function for its intended lifetime. This element includes review/oversight of product assurance efforts of partners and subcontractors for compliance with the SoloHI PAIP.
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Responsible WBS WBS Task Name and Description Manager / No. Organization Science / Technology: This summary element includes the managing, directing, and controlling of the science investigation aspects, as well as leading, managing, and performing the technology demonstration elements of the Project. The costs incurred to cover the Principal Investigator, Project Scientist, science team members, and equivalent personnel for technology 4.0 demonstrations are included. Specific responsibilities include defining the PI/NRL science or demonstration requirements; ensuring the integration of these requirements with the payloads, spacecraft, ground systems, and mission operations; providing the algorithms for data processing and analyses; and performing data analysis and archiving. This element excludes hardware and software for onboard science investigative instruments/payloads. NRL Science Team: Includes all NRL science team activities as defined in 4.0 4.1 PI/NRL and travel JPL Science Team: Includes all JPL science team activities as defined in 4.0 4.2 Co-I/JPL and travel SoloHI Science Payload Development and Test: This element includes the equipment provided for special purposes in addition to the normal equipment (i.e., GSE) integral to the SoloHI instrument. This includes leading, managing, 5.0 PE/NRL and implementing the hardware and software payloads that perform the scientific experimental and data gathering functions placed on board the spacecraft. 5.1 Mechanical Systems IM/NRL Structures: The equipment, data, services, human resources, and facilities required to design and fabricate the structural elements of the SoloHI. This 5.1.1 IM/NRL element also includes all mass simulators and/or structural models of all the SoloHI systems. Door Deployment Mechanism: The equipment, data, services, human resources, and facilities required to develop the one-shot door deployment 5.1.2 IM/NRL mechanism for the SoloHI. Includes the design, fabrication, integration; deployment testing; and subsystem environmental testing. Thermal Design/Components: The equipment, data, services, human resources, and facilities required to develop the thermal control system for the 5.1.3 IM/NRL SoloHI. Includes the design, fabrication, integration, test, MLI, and any thermal control hardware (thermistors, heaters, etc.) Instrument Mounts: The equipment, data, services, human resources, and 5.1.4 facilities required to develop the SoloHI to spacecraft mounts. Includes the IM/NRL design, fabrication, integration, test, and hardware. 5.2 Electronics Systems NRL SoloHI Control Electronics (SCE): The equipment, data, services, human resources, and facilities required to develop the elements of the SoloHI control 5.2.1 NRL electronics. Includes the design, fabrication, integration, test, structure, Engineering Development Model (EDM), and EEE parts.
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Responsible WBS WBS Task Name and Description Manager / No. Organization Camera Electronics: The equipment, data, services, human resources, and facilities required to develop the elements of the SoloHI camera electronics. 5.2.2 NRL Includes the design, fabrication, integration, test, structure, Engineering Development Model (EDM), and EEE parts. SoloHI Harness: The equipment, data, services, human resources, and 5.2.3 facilities required to develop the elements of the SoloHI Harness. Includes the NRL design, fabrication, integration, test, and EEE parts. 5.3 Focal Plane Array Systems NRL FPA Structures: The equipment, data, services, human resources, and facilities required to develop the FPA assembly for the SoloHI. Includes the 5.3.1 NRL design; fabrication; itegration with the APS detector system; and subsystem testing FPA Detector: The equipment, data, services, human resources, and facilities 5.3.2 required to develop the APS detector for the SoloHI. Includes the design, NRL fabrication, integration, and testing. FPA Optics: The equipment, data, services, human resources, and facilities 5.3.3 NRL required to design and fabricate the optical elements of SoloHI. 5.4 Software SWE/NRL Payload Flight Software: The human resources, equipment, data, services, and facilities required to develop, produce, test and deliver the Flight Payload Software System to meet science, operations, and data requirements. The 5.4.1 Flight Software includes: health monitoring; data storage and handling; and SWE/NRL software required to interface with command and to control the payload. Includes support of I&T, system test and launch preparations; and support of operations planning and operational readiness verification and validation. Payload Ground Software: The human resources, equipment, data services 5.4.2 and facilities required to develop, produce, test and deliver the Ground SWE/NRL Software System to meet system requirements 5.5 SoloHI Instrument Assebly, Integration and Test (AI&T) PE/NRL SoloHI Instrument Integration: The human resources, equipment, data, 5.5.1 services, and facilities required to integrate and functionally test the SoloHI PE/NRL instrument. SoloHI Instrument Environmental Testing: The human resources, equipment, 5.5.2 data services and facilities required to environmentally test the SoloHI PE/NRL instrument. SoloHI Calibration: The human resources, equipment, data services and 5.5.3 facilities required to calibrate the SoloHI instrument and characterizethe SoloHI PI/NRL payload performance. 6.0 Spacecraft (S/C) Development (Not Applicable) 7.0 Mission Operations and Data Analysis MOM/NRL
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Responsible WBS WBS Task Name and Description Manager / No. Organization Payload Operations: Includes the workforce and services required to assemble, integrate, train, and test personnel to coordinate SoloHI payload operations with the Solar Orbiter Mission Operations Center. Also includes all labor, subcontracts, materials and other direct costs to define the Science 7.1 MOM/NRL Operations Center located at the NRL. This includes detailed design, hardware procurement, system integration, operational procedures development, integration testing, and simulation support for the Solar Orbiter Mission Operations Center. NRL Data Analysis: At NRL, includes all labor, subcontracts, materials and other direct costs to develop hardware, software, and operational processes to perform science data processing including raw data ingest, frame 7.2 PI/NRL synchronization, data decoding, error detection and correction, data sorting, data management, data reprocessing, etc. Includes any algorithms necessary to process and analyze the measurements taken from the SoloHI instrument. JPL Data Analysis: At JPL, includes all labor, subcontracts, materials and other direct costs to develop hardware, software, and operational processes to perform science data processing including raw data ingest, frame 7.3 CO-I/JPL synchronization, data decoding, error detection and correction, data sorting, data management, data reprocessing, etc. Includes any algorithms necessary to process and analyze the measurements taken from the SoloHI instrument. 8.0 Launch Vehicle and Services (Not Applicable) 9.0 Ground Tracking Network (Not Applicable) Solar Orbiter Systems Integration and Testing: This element includes the H/W, S/W, procedures, and personnel to support the SoloHI integration to the Solar 10.0 PE/NRL Orbiter S/C. Also includes support for environmental testing; launch processing; and on-orbit checkout. Systems I&T: This element includes the H/W, S/W, procedures, and personnel 10.1 to support the SoloHI integration to the Solar Orbiter S/C. Also includes PE/NRL support for environmental testing. Systems Launch Processing: This element includes the H/W, S/W, 10.2 procedures, and personnel to support, as needed, launch processing of the PE/NRL Solar Orbiter Observatory. Systems On-Orbit Checkout: This element includes the H/W, S/W, procedures, 10.3 and personnel to support, as needed,on-orbit checkout of the SoloHI PE/NRL instrument on the Solar Orbiter Observatory. Education and Public Outreach (E/PO): This summary element supports the responsibilities of NASA’s missions, projects, and programs in alignment with 11.0 the Strategic plan for Education (includes Mgmt and coordinated activities, PI/NRL formal education, informal education, public outreach, media support, and web site development).
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Appendix B. TRADE STUDIES SUMMARY Table B-1. Trade Studies Summary
Trade Study Comparison Metrics Results Status Title Flight Spacewire Interface NRL will use the LEON3 FPGA as the flight Closed Processor Compatibility, processor. It satisfies the SoloHI processing Selection Availability, Heritage, requirements with adequate margin, is (LEON2 ASIC Minimize Electronics immediately available, and has flight heritage vs. LEON3 Development, Minimize (LEON3 on Chandrayaan-1, RTAX2000S on FPGA) Board Area many missions). The LEON3 FPGA design allows the inclusion of a Memory Controller, and Spacewire interface that would require additional electronics to be added to the board design if the LEON2 ASIC had been chosen. The Memory Controller and Spacewire modules for the Gaisler Research LEON3 have been developed already. APS Detector Full Well Capacity, NRL will use the 6-transistor (6T) pixel design Closed Pixel Design Linearity, Radiation for the SoloHI APS/CMOS detector. The 6T (3T, 5T or 6T) Tolerance, Read Noise pixel is the same as the 5T pixel with an extra transistor that increases the capacitance at the sense node and increases the full well. This extra transistor also allows the detector to switch between high and low camera gains, which will allow the SoloHI detector to reduce the read noise for the faint signal regions in the outer FOV. The 5T and 6T linearity is better than the 3T performance. The radiation tolerance for all 3 pixel designs is expected to be more than adequate.. SCE Real-time Satisfies OS NRL will use the free, open-source Real-Time Closed Operating requirements, Staff Executive for Multiprocessor Systems (RTEMS) System learning curve, Price, as the real-time operating system for the SoloHI (RTEMS, Support Costs Control Electronics Box. It satisfies all SoloHI VxWorks 5.4, OS requirements. Hammers has developed VxWorks 6.0) software on RTEMS for Themis. Straylight Straylight Rejection at We studied a variety of surface finishes and side Closed Rejection Side Aperture from S/C wall baffle configuration. We selected the Baffle/ structures, Mass, AIT simplest design which gave reasonable Side Wall complexity performance. Candidates
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Trade Study Comparison Metrics Results Status Title SIM Placement Scientific return from SoloHI observes from about 5-45° from the sun- Further on Spacecraft collaborative spacecraft line, so, looking at the west limb on Research in for Science instrumentation and the ram side, it would remotely observe CMEs Phase B Observations spacecraft. and quiescent solar structures such as Planned (RAM-side vs. streamers before the in-situ instruments observe Anti-RAM side) them. From the anti-ram side, it would still see CMEs in the same region, but not see the same ones that the spacecraft observes in-situ. Opportunities for viewing the Solar Probe Plus are about the same from either side. The most significant observation is the remote detection of the source of energetic particles, which can only be with SoloHI on the ram side. A presentation of the study was given to the Solar Orbiter SWT in June 2010, which recommended that the spacecraft team revisit the accommodation of the instruments with the goal of placing SoloHI on the ram side. Evaluation of Measured BRDF Two candidate black anodize processes were Closed. black baffle performance at NRL identified which gave suitable performance. The coatings first “Deep Space Black” is provided by N- Sciences. The second “laser-black” is provided by Epner. Coupons were procured from both companies. Epner provided coupons. N- sciences has not yet provided a coupon. Evaluation of both coatings were made at NRL and the performance was comparable. The laser black coating does give acceptable performance and was selected. overall resource requirements, The spacecraft accommodation of the SCE was Closed. electronics testing and verification changed to co-locate the box near the SIM. The architecture approach, cable SoloHI team identified an optimized electronics length/signal drive architecture for this location. The optimized considerations.. architecture consolidates the processor card and the camera electronics card digital elements. This consolidated digital card is located on the SIM.
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Appendix C. PHASE B WORK AGREEMENTS The majority of NRL support contractors for SoloHI are pulled from existing Omni-bus contracts. The contract for the SoloHI APS detectors is in place. The only other contract needed for SoloHI will be the contract for the Optics. This will be placed once the design is settled during Phase B.
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Appendix D. PRODUCT ASSURANCE IMPLEMENTATION PLAN Reference the SoloHI Product Assurance Implementation Plan, SSD-PLN-SOLOHI-0002, delivered separately.
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Appendix E. SoloHI DETECTOR REQUIREMENTS, SELECTION, AND BACKUP The essential SoloHI requirement of recording wide field coronal images over a large dynamic range at reasonable cadence drives the SoloHI detector scientific requirements. The SoloHI programmatic resources of mass and power as well as cost/complexity and radiation tolerance considerations were also considered for selecting the SoloHI detector. E.1 SoloHI Detector Driving Requirements Table E-1 shows the SoloHI detector driving requirements and the associated top-level science and instrument requirements. This set of detector driving requirements meet the SoloHI field of view, resolution, photometric accuracy and image cadence requirements. The detector radiation hardness meets the SoloHI performance requirements at the end of the Extended Science mission, also known as End of Life (EOL). This capability allows the detector to meet the strict SoloHI accommodations requirements by minimizing the required shielding. A detector that does not have a high radiation tolerance could still be selected by increasing the shielding to minimize the radiation dose at the detector. Table E-1. Detector Driving Requirements that Led to the APS Detector Selection
The SoloHI detector requirements for the imaging area, quantum efficiency, linear full well capacity, read noise, dark current, and charge transfer efficiency are derived as a set of
E-1 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 requirements which satisfy the SoloHI photometric accuracy at EOL. This is not a unique solution to this trade space and our trade study considered a solution set tailored to take advantage of the strengths of each proposed detector technology. The solution sets examined for both the APS, CCD and hybrid technology satisfy the top-level SoloHI photometric accuracy and image cadence requirements with positive margin. E.2 SoloHI Detector Trade Study Summary Many detectors were considered in the initial trade study to select the SoloHI detector. Emerging detector technologies both in the US and Europe were considered as possible candidates. A parallel detector development for Solar Orbiter is being implemented by the EUI and PHI teams. This represents an additional possible backup. We identified three detectors as viable candidates. These detectors satisfy the SoloHI functional and performance requirements. Table E-2 summarizes their design, performance and technical maturity characteristics. These detectors include an APS candidate, a hybrid detector candidate and a CCD candidate. Table E-3 describes the programmatic and technical accommodation for these detectors. Table E-2. SoloHI Detector Candidate Characteristics Detector Hawaii 4RG 4Kx4K Minimal CCD – 231 Characteristic Manufacturer Teledyne Imaging Systems Sarnoff Corporation e2v Type Hybrid array, 100 micron Hy- front side illuminated, Back-side illuminated Visi Si, bump bonded to a read 6T active pixel Charge Coupled out integrated circuit sensor, stitched Detector design Analog to digital off-chip off-chip off-chip converter Readout 64 output ports available, 2-16 output ports 4 ports, one at each architecture interleaved guide window available, transferred corner bidirectional readout available by row, 1Kx1K block serial transfer in 1K row enable blocks to serial register Array Size 4096 x 4096 >4000x4000 4096x4112 Pixel Size 10 microns 10 microns 15 microns Pixel to Pixel <4% standard deviation ~3% standard <<1% uniformity deviation Linearity <2% <2%, still being <1% tested with the SoloHI minimal Linear Full Well ~80,000 electrons >80,000 350,000 typ. Read noise 10-15 electrons (rms) for A2 <2 electrons (rms) at <10 electrons (rms) ROIC high gain, <30 electrons low gain. Maximum 5 MHz using multiple ports 4 MHz/port 3 MHz/port Readout Rate Dark Current <1 electron/s/pixel at <200 K <5 electron/s/pixel at <1 electron/s/pixel at <240 K 210 K Cosmetics 99.6% of pixels at >85% of Goal 99.8% pixels 99.9% expected QE scientifically usable pixels.
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Detector Hawaii 4RG 4Kx4K Minimal CCD – 231 Characteristic Radiation tested to 90 krad with gamma Tested to 130krad design similar to 42 Tolerance radiation 137Cs, tested to 5krad with 63MeV protons, family, devices perform with protons >1Mrad with gamma well if cooled to <190K radiation Operating <193 K < 240 K <193 K Temperature Imaging pixel limited with some cross Pixel limited pixel limited Resolution talk Quantum >80% ~ 25% >70% Efficiency over 500 to 700 nm Technology 5 – 6 5 7 Readiness Level radiation testing results need extensive laboratory device family further review, may require testing and radiation extensively tested and more testing or more shielding testing on prototype characterized if device performance has not devices been verified at sufficient dose.
Table E-3. Detector Candidate Programmatic and Technical Accommodation Requirements Accommodation Hawaii 4RG 4Kx4K Minimal CCD-231 Thermal larger radiator required to Baseline Larger radiator reach 190 K required to reach 190 K Shielding TBD, limited proton testing, 3 mm, TBR after <12 mm would expect current shielding Sandbox VI radiation to be adequate. tests Electronics Feedback from SAO and NRL on-chip ACDS, simple Typical for a CCD implementation groups working with this readout, requires digital requiring control of detector indicate that the control of the detector, rail voltages, ACDS readout implementation is conditioning of Vdd and with ADC off-chip. complex. buffering/ADC of the signal. Shutter Not required not required required for full frame readout, could be operated as a frame transfer device for half the area. Mass CBE Delta Similar to Baseline Baseline <2 kg (kg) Avg Power CBE Similar to Baseline Baseline ~4 W Delta (W) Detector Cost $2.0M $2.5M $0.9M Estimate Camera $3.1M, may be able to exploit $1.1M, current CEBB $2.1M Electronics Cost program efforts to reduce drive electronics shows (incl shutter and Estimate costs. that the camera design shutter electronics) will be straightforward.
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Accommodation Hawaii 4RG 4Kx4K Minimal CCD-231 Comments device design still undergoing device design under identical type of refinement, perceived high refinement and test, detector previously complexity of array and Residual fabrication and flown on HI associated readout design risk to be retired. instrument, device electronics, may be able to exists with well exploit existing program efforts defined interfaces and to reduce cost. drive electronics. The Sarnoff 4k x 4k APS detector was selected as the baseline. The Sarnoff APS satisfies the SoloHI functional, performance, and resource allocation requirements at the lowest overall development risk. The Sarnoff APS detector development program is focused primarily on scaling the minimal design to meet the pixel size, dynamic range, readout speed and array size requirements. Preliminary results from the detector laboratory evaluation show that the SoloHI minimal prototype detector meets the program requirements. These results are summarized in section 3 of the updated concept study report. Engineering work on the SoloHI flight detector is well advanced with a PDR scheduled for January 2012 and a CDR in the spring of 2012. The e2v CCD detector was also evaluated as a possible option. Among the three options, the CCD option has the lowest overall cost and development risk. The detectors can be readily procured. The camera and shutter electronics development is straight-forward with measured performance on previous electronics that meets or exceed the SoloHI requirements. However, this option was not selected as the SoloHI detector, because it requires additional mass and power above the SoloHI resource allocation. The principal scientific disadvantage of the CCD is that the rapid cadence small turbulence mode (requiring windowed readout) may be slightly compromised. Accommodation of the required shutter based requires additional electronics and software. Additional shielding to accommodate the less robust CCD technology will likely be required as shown on the trade table. The Teledyne Imaging Hawaii detector satisfied the SoloHI functional, performance, and resource allocation requirements, but at the highest cost of any detector candidate and at the highest development risk for the combined detector and camera electronics system. Although the Hawaii detector option does not require significant detector development, it does require extensive camera development. The operational complexity of the existing Hawaii array and its Sidecar ASIC was viewed as a significant development risk. The SoloHI team has based this assessment of development risk on interviews with other NRL and Smithsonian Astrophysics Observatory (SAO) design teams about their efforts to develop camera electronics for the Teledyne Imaging Hawaii detector and its Sidecar ASIC for other applications. The primary use of the Hawaii detector for JWST and ground based instrumentation has been for a benign radiation environment and the array will likely need further testing to qualify it for the SoloHI environment. All of the SoloHI detector candidates (even the off-the-shelf detectors) would require further evaluation to demonstrate that they can survive and meet the performance requirements at the baseline operating temperature and shielding. E.3 SoloHI Detector Backup Plan Figure E-1 shows the contingency plans and the related trigger dates for the SoloHI APS detector development program. There are 3 SoloHI detector contingency plans that will require a different detector than the baseline. If there is a large functional or performance requirement violation in the ongoing Sandbox VI radiation/performance testing or a major performance issue which
E-4 Version 2.0 – 31 December 2011 SoloHI Concept Study Report SSD-RPT-SOLOHI-0003 surfaces at the upcoming PDR, the SoloHI team will need to baseline a different detector from another vendor to satisfy the SoloHI requirements and to have a detector with a technology readiness level greater than 6 by the SoloHI PDR. Both the Teledyne Imaging Hawaii 4RG and the e2V 231-84 CCD would satisfy the SoloHI functional and performance requirements. However, the primary SoloHI backup detector for the Sandbox VI test failure is the Teledyne Imaging Hawaii 4RG, because it will remain within the current SoloHI mass and average power allocations, maintain the appropriate resource margins for Phase A, and the cost of this contingency plan can be covered with the existing SoloHI cost reserves. The SoloHI team still considers the e2v CCD to be attractive due to its lower costs and limited development requirements. It could be implemented in the current SoloHI mass allocation with a NASA waiver for the mass margin at Phase A, but it would exceed the current SoloHI average power allocation. Therefore, the SoloHI team will continue discussions with ESA to determine if the SoloHI power allocation could be increased to make this a viable backup option. The average power allocation delta could be minimized by exploring instrument descopes to the SoloHI baseline science, such as reducing the detector array size from 4k x 4k to 2k x 2k. The other SoloHI detector contingency plans are based on a failure in the SoloHI flight detector program. If the flight detector fabrication run does not produce detectors that satisfy the SoloHI functional/performance requirements, the SoloHI team and Sarnoff will investigate the cause of failure. A second foundry run of the SoloHI flight detector will be performed if the failure can be addressed by a design or fabrication process change. If the failure is related to the stitching of the detector blocks, the SoloHI team will design and fabricate a monolithic APS detector with the same pixel design and high speed readout circuitry as the baseline, but with a smaller array size. Table E-4 summarizes the impact of each of these detector contingency plans. For all of these contingency plans, the budget impact ranges from $800K for another Jazz fabrication run to $1.5M additional cost for the hybrid option. The allowed procurement time frame for the backup CCD or hybrid array option is 3 months for an engineer grade detector and <12 months for the flight model detectors. This should be sufficient time to delivery the backup detectors. With a final decision in April 2012, flight detectors would be available at the latest in April 2013. The SoloHI instrument master schedule incorporates a date of April 2013 date for the Sarnoff detector delivery date. This date allows for some margin on the current projected Sarnoff delivery date of February 2013. For the CCD and hybrid array options, the SoloHI team will procure engineering model detectors for testing with the flight electronics. Each of these contingency plans can be successfully executed within existing SoloHI resources and will not delay the flight instrument delivery for observatory integration.
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Large violation Select backup Technology detector Fabricate SB VI Test SB VI activities N devices devices successful? March 2012
Generate new March 2012 error budget Y Small violation
Finish FM March 2012 Fabricate 4K × detailed design 4K devices activities
Select baseline 4K Fabrication Y × 4K devices NRL successful? receive September 2012 February 2013
N
4K × 4K devices Process (P) or stitching Second run to delivery P fabricate 4K× 4K (S) problem? May 2013 devices
monolithic S Fabricate 2K × 2K backup monolithic 2K × 2K devices delivery May 2013
Figure E-1. SoloHI Detector Contingency Plans and Decision Tree. Dates in the figure are based on the current Sarnoff schedule. Detector delivery date in the SoloHI instrument master schedule is April 2013. Delivery date for backup flight detectors (e2v CCD or hybrid) is also April 2013 with a 12 month procurement cycle. Engineering CCD and hybrid backup detectors would be available in June 2012 Table E-4. Projected Budget Impact for SoloHI Detector Contingency Plans Contingency SoloHI Budget Contingency Plan Description Plan Number Delta 1 Procure Teledyne Imaging Hawaii detectors $1.75 M 2 2nd Fabrication Run of Sarnoff Minimal APS 4k x 4k detector $0.8 M 3 Monolithic Sarnoff Minimal APS 2k x 2k Detector $1.2 M Plan 1 includes additional $250K for Sarnoff development costs for decision made in April 2012.
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Appendix F. CONTAMINATION CONTROL PLAN Reference the SoloHI Contamination Control Requirements and Plan, SSD-RQT-SOLOHI-0004, dated 16 November 2010.
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Gosling, J. T. & Riley, P. (1996) The acceleration of slow coronal mass ejections in the high speed solar wind, Geoph. Res. Lett., 23, 2867. Hayes, A. et al. (2001) Deriving the Electron Density of the Solar Corona from the Inversion of Total Brightness Measurements. Astroph. J.,548, 1081. Ho, G. C. et al. (2000) Enhanced solar wind 3He2+associated with coronal mass ejections, Geophys. Res. Lett., 27, 309. Hundhausen, A. J. (1993) Sizes and locations of coronal mass ejections -- SMM observations from 1980 and 1984-1989, J. Geophys. Res., 97,1619. Illing, R. M. E. & Hundhausen, A. J. (1983) Disruption of a coronal streamer by an eruptive prominence and coronal mass ejection, J. Geophys. Res., 91, 10951. Jackson, B. V. et al. (1985) HELIOS spacecraft and Earth perspective observations of three looplike solar mass ejection transients, J. Geophys. Res.,90, 5075. Jackson, B. V. & Leinert, C. (1986) HELIOS images of coronal mass ejections, J. Geophys. Res.,90, 10759. Kahler, S. W. (2001), The correlation between solar energetic particle peak intensities and speeds of coronal mass ejections: Effects of ambient particle intensities and energy spectra, J. Geophys. Res., 106, 20,947-20,955. Lee, M. A., (2000) Acceleration of Energetic Particles on the Sun, in the Heliosphere and in the Galaxy, in AIP Conf. Ser. 528, ed. R.A. Mewaldt et al. (Melville: AIP), 3. Leinert, C. et al. (1998) The 1997 Reference of Diffuse Night Sky Brightness, A&A Supp., 127,1. Low, B. C. (1982) Self-similar MHD: II. The expansion of a stellar envelope into a surrounding vacuum. Mann, I., Krivov, A., and Kimura, H. (2000) Dust Cloud Near the Sun, Icarus, 146, 568. Manoharan, P. K. et al. (2001) Coronal mass ejection of 2000 July 14 flare event: Imaging from the near-Sun to Earth environment, Astroph. J, 559, 1180. Mason, G. M., et al. (1999) 3He enhancements in large solar energetic particle events, Astrophys. J., 525, L133. Mewaldt, R. A. et al. (2001) How Efficient are Coronal Mass Ejections at Accelerating Solar Energetic Particles?, Proc. of Solar Wind 11,Whistler Canada. Mulligan, T et al. (1999) Intercomparison of NEAR and Wind interplanetary coronal mass ejection observations, J. Geophys. Res., 104,28217. Newmark, J. S., et al. (2003) 3D Electron Density Reconstruction from the SECCHI White Light Coronagraphs Onboard Stereo, Bulletin of the AAS, 35, 807. Odstrcil, D. & Pizzo, V. J. (1999a) Three-dimensional propagation of CMEs in the structured solar wind flow: 1. CME launched within the streamer belt., J. Geophys. Res., 104, 483. Odstrcil, D. & Pizzo, V. J. (1999b) Three-dimensional propagation of CMEs in the structured solar wind flow: 2. CME launched adjacent to the streamer belt., J. Geophys. Res., 104, 493. Odstrcil, D. et al. (2004) Numerical simulation of the 12 May 1997 interplanetary CME event, J. Geophys. Res., 109, A02116.
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Odstrcil, D. et al. (2005) Propagation of the 12May 1997 interplanetary coronal mass ejection in evolving solar wind structures, J. Geophys. Res., 110, A02106. Odstrcil, D. & Pizzo, V. J. (1999c) Distortion of the interplanetary field by three-dimensional propagation of coronal mass ejections in a structured solar wind, J. Geophys. Res., 104, 28225. Ontiveros, V., Vourlidas, A. (2008), Quantitative Measurements of CME-driven Shocks from LASCO Observations, ApJ, submitted. Ontiveros, V., Vourlidas A. (2007) How do CME-Shocks Look Like?: Study of Shock Geometry. Eos Trans. AGU, 88(52), Fall Meet. Suppl., Abstract SH31A-0223. Plunkett, S.P. et al. (2000) Simultaneous SOHO and ground-based observations of a large eruptive prominence and coronal mass ejection, Sol. Phys., 194, 371. Ragot, B. R., Kahler, S. W. (2003) Interactions of Dust Grains with Coronal Mass Ejections and Solar Cycle Variations of the F-Coronal Brightness, Astrophysical Journal 594, 1049- 1059. Rakowski, C. E. et al. (2007) Ion Charge States in halo coronal mass ejections: What can we learn about the eruption?, ApJ, 667, 602. Reames, D. V. (1999), Particle acceleration at the Sun and in the heliosphere, Space Sci. Rev., 90,413. Richardson, I. et al. (2003) Are CME „interactions‟ really important for accelerating major solar energetic particle events?, Geophys. Rev. Lett., 30,SEP 2-1. Riley, P. et al .2003) Dynamical evolution of the inner heliosphere approaching solar activity maximum: interpreting Ulysses observations using a global MHD model, Ann. Geophys., 21,1347. Riley, P. et al. (2004) Ulysses observations of the magnetic connectivity between coronal mass ejections and the sun, ApJ, 608, 1100. Sheeley, N. R., Jr., et al. (1997) Measurements of flow speeds in the corona between 2 and 30 Rs, Astrophs J., 484, 472. Sheeley, N. R. et al. (1999) Continuous tracking of coronal outflows: Two kinds of coronal mass ejections, J. Geophys. Res., 104, 4739. Sheeley, N. R. Jr., & Wang, Y.-M. (2001) Coronal inflows and sector magnetism, Astrophys. J., 562, L107. Sheeley, N. R. Jr et al. (2008) Heliospheric Images of the Solar Wind at Earth, Astrophys. J., 675 (Mar. 1,2008). Skoug, R. M. et al. (1999) A prolonged He+ enhancement within a coronal mass ejection in the solar wind, Geophys. Res. Lett., 26, 161. Smith, E. J. & Balogh, A., (2003) Open Magnetic Flux: Variation with Latitude and Solar Cycle, SOLAR WIND TEN: Proceedings of the Tenth International Solar Wind Conference; DOI: 10.1063/1.1618543, AIP Conf. Proc. Volume 679, pp. 67-70. Suess, S. T., et al. (1996) Volumetric heating in coronal streamers, J. Geophys. Res., 101, 19957. Thernisien, A. F.; & Howard, R. A. (2006) Electron Density Modeling of a Streamer Using LASCO Data of 2004 January and February, Astrophys. J., 642, 523. Thernisien, A. F et al. (2006) Modeling of Flux Rope Coronal Mass Ejections, ApJ, 652, 763.
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Thernisien, A. F et al. (2007) Forward modeling of reconstruction techniques applied to STEREO/SECCHI data, AGU Fall Mtg, abstract SH32A0778. Tylka, A. (2005) Shock Geometry, Seed Populations, and the Origin of Variable Elemental Composition at High Energies in Large Gradual Solar Particle Events, Astroph. J., 625, 474. Vourlidas, A. et al. (2000) LASCO measurements of the energetics of coronal mass ejections, ApJ,534, 456. Vourlidas, A. et al. (2003) Direct Detection of a Coronal Mass Ejection-Associated Shock in Large Angle and Spectrometric Coronagraph Experiment White-Light Images, Astroph. J., 598, 1392. Vourlidas, A. & Howard, R. A. (2006) The proper treatment of coronal mass ejection brightness: A new methodology and implications for observations, ApJ, 642, 1216. Vourlidas, A. et al. (2007) First direct observation of the interaction between a comet and a coronal mass ejection leading to a complete plasma tail disconnection, Astrophys. J. L, 668, 79. Vourlidas, A., & Riley, P. (2007) Direct Imaging of the Heliospheric Plasma Sheet from the SECCHI telescopes on the STEREO Mission, Eos Trans. AGU, 88(52), Fall Meet. Suppl., Abstract SH21A-0283. Vrsnak, B. & Gopalswamy, N. (2002) Influence of the aerodynamic drag on the motion of interplanetary ejecta, J. Geophys. Res., 107, 1019. Wang, Y.-M. et al. (1999a) Coronagraph observations of inflows during high solar activity, GRL 26, 1203. Wang, Y.-M. et al. (1999b) Streamer disconnection events observed with the LASCO coronagraph, GRL 26, 1349. Wang, Y.-M., et al. (2000) The long-term variation of the Sun‟s open magnetic flux, GRL, 27, 4, 505. Wang, Y.-M. & Sheeley, N. R. Jr (2006) Observations of Flux Rope Formation in the Outer Corona, ApJ, 644, 638. Webb, D. F. & Jackson, B. V. (1990) The identification and characteristics of solar mass ejections observed in the heliosphere by the HELIOS-2photometers, J. Geophys. Res., 95, 20641. Webb, D. F. et al. (2006) Solar Mass Ejection Imager (SMEI) observations of coronal mass ejections in the heliosphere, J. Geophys. Res., 111, A12101. Wu, C. C. & Lepping, R. (2006) Characteristics of magnetic cloud/magnetic-cloud-like structures during the years 1995-2003, 36th COSPAR Assembly, 36, 40. Yashiro, S. et al. (2004) A catalog of white light coronal mass ejections observed by the SOHO spacecraft, J. Geophys. Res., 109, A7, CiteIDA07105. Zurbuchen, T. H. & Richardson, I. G. (2006) In-situ solar wind and magnetic field signatures of interplanetary coronal mass ejections, Sp. Sci. Rev., 123, 31.
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Appendix H. LIST OF ACRONYMS AA/SMD Associate Administrator for the Science Mission Directorate AFT Abbreviated Functional Test AO Announcement of Opportunity APS Active Pixel Sensor ASIC Application Specific Integrated Circuit BOE Basis of Estimate CADRe Cost Analysis Data Requirement CASB Cost Accounting Standards Board CCB Configuration Control Board CCD Charge Coupled device CCMC Community Coordinated Modeling Center CCRP Contamination Control Requirements and Plan CCSDS Consultative Committee for Space Data Systems CDRL Contract Data Requirements List Program CER Cost Estimating Relationship CFRP Carbon Fiber Reinforced Plastic CIR Co-Rotating Interaction Region CME Coronal Mass Ejection CMP Configuration Management Plan COTS Commercial Off-the-Shelf CPT Comprehensive Performance Test CR Confirmation Review CRM Continuous Risk Management CSG Centre Spatial Guyanais CTE Charge-Transfer-Efficiency CVCM Collected Volatile Condensable Material DFAR Defense Federal Acquisition Regulation DSM Deep Space Maneuvers E/PO Education and Public Outreach EAR Export Administration Regulations EDAC Error Detection and Correction EDM Engineering Development Model EED Electro Explosive Device EEE Electrical, Electronic, and Electromechanical EELV Evolved Expendable Launch Vehicle EGSE Electrical Ground Support Equipment EID Experiment Interface Document EM Engineering Model EMC Electromagnetic Compatibility EMI Electromagnetic Interference ENM End of Nominal Mission ESA European Space Agency EVM Earned Value Management EVMS Earned Value Management System FAR Federal Acquisition Regulation
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FEA Finite Element Analysis FITS Flexible Image Transport System FMEA Failure Modes and Effects Analysis FOR Flight Operations Review FOSO Focused Opportunity for Solar Orbiter FOT Flight Operations Team FOV Field of View FPA Focal Plane Assembly FRR Flight Readiness Review FSW Flight Software FT Functional Tests FTA Fault Tree Analysis G&A General and Administrative GAM Gravity Assist Maneuver GDS Ground Data System GEVS General Environmental Verification Specification GIDEP Government-Industry Data Exchange GPR Goddard Procedural Requirements G-RBSP GEOSPACE-Radiation Belt Storm Probes GSE Ground Support Equipment GSFC Goddard Space Flight Center GUI Graphical User Interface HCS Heliospheric Current Sheet HELEX Heliophysical Explorers HI Heliospheric Imager HPS Heliospheric Plasma Sheet HSPD Homeland Security Presidential Directive H/W Hardware I/F Interface I&T Integration and Test IACO Integration, Assembly, and Checkout IAT Integration, Assembly, and Test ICD Interface Control Document ICE Independent Cost Estimates ICME Interplanetary CME ICR Initial Confirmation Review ICWG Integration Working Group Meetings IDL Interactive Data Language IMS Integrated Master Schedule IPT Integrated Product Team ITAR International Traffic in Arms Regulations JPL Jet Propulsion Laboratory JSTDT Joint Science and Technology Definition Team KDP Key Decision Points KPP Key Performance Parameters KSC Kennedy Space Center
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LBA Lens Barrel Assembly LEOP Launch and Early Operations LLIS Lessons Learned Information System LOA Letter of Agreement LOE Level of Effort LRR Launch Readiness Review LRU Line Replaceable Unit LVDS Low Voltage Differential Signaling M&P Materials and Processes MAE Mission Assurance Engineer MAHRSI Middle High Resolution Spectrograph Investigation MAP Mission Archiving Plan MAR Mission Assurance Requirements MEL Master Equipment List MGSE Mechanical Ground Support Equipment MHD Magneto-hydro-dynamical MIP Mission Implementation Plan MIPS Million Instructions Per Second MO&DA Mission Operations and Data Analysis MOC Mission Operations Center MOM Mission Operations Manager MOR Mission Operations Review MoU Memorandum of Understanding MOWG Mission Operations Working Group MRR Mission Readiness Review MRT&V Mission Requirements, Traceability, and Verification NAR Non-Advocate Review NASA National Aeronautics and Space Administration NICM NASA Instrument Cost Model NID NASA Interim Directive NIEL Non-Ionizing Energy Loss NPPL NASA Preferred Parts List NRL Naval Research Laboratory NWCF Navy Working Capital Fund OBS Organizational Breakdown Structure OH Overhead PAIP Product Assurance Implementation Plan PCB Parts Control Board PCC Processor Control Card PCP Parts Control Plan PDD Payload Definition Document PDMP Project Data Management Plan PDR Preliminary Design Review PED Plastic Encapsulated Device PER Pre-Environmental Review PI Principal Investigator
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PM Project Manager PMB Performance Measurement Baseline PMCS Project Management and Control System PMO Project Management Office POC Payload Operations Center PRA Probability Risk Assessment PSR Pre-Ship Review QA Quality Assurance QAE Quality Assurance Engineer QMS Quality Management System RDM Radiation Design Margin RE Reliability Engineer RGA Residual Gas Analyzer RHA Radiation Hardness Assurance RMP Risk Management Plan RTC Remote Terminal Control S/C Spacecraft SB Small Business SCE SoloHI Control Electronics SDB Small Disadvantaged Business SDO Solar Dynamics Observatory SE System Engineer SEE Single Event Effect SEL Single Event Latchup SEMP Systems Engineering Management Plan SEP Solar Energetic Particle SEU Single Event Upset SIM SoloHI Instrument Module SMD Science Mission Directorate SMEI Solar Mass Ejection Imager SMEX Small Explorer SMSR Safety and Mission Success Review SOC Science Operations Center SODA Science Operations and Data Analysis SOHO Solar and Heliospheric Observatory SoloHI Solar Orbiter Heliospheric Imager SOW Statement of Work SQAP Software Quality Assurance Plan SRD System Requirements Document SRR Systems Requirements Review SRTM Science Requirements Traceability Matrix SSD Space Science Division SSR&MA System Safety, Reliability, and Mission Assurance STEREO Solar Terrestrial Relations Observatory TID Total Ionizing Dose TMC Total Mission Cost
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TML Total Mass Loss TOPS Top Teachers of Physical Science TPM Technical Performance Measurement TQCM Thermoelectric Quartz Crystal Microbalance TRL Technology Readiness Level TVAC Thermal Vacuum V&V Verification and Validation VSO Virtual Solar Observatory WBS Work Breakdown Structure WWW World Wide Web
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Appendix I. INTEGRATED MASTER SCHEDULE SoloHI Integrated Master Schedule is provided in separate Project and PDF files.
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Appendix J. MASTER EQUIPMENT LIST (MEL) Table J-1 presents the Master Equipment List (MEL) for the SoloHI science payload. The mass contingency percentage for each SoloHI subsystem is based on the maturity of the basis of mass estimate (BME). Table J-1 defines the basis of estimate codes used in the MEL. The SoloHI mass components that have an “Estimate” mass maturity include thermal components, secondary structure, and fasteners. The Science Payload thermal component mass is estimated as 3.0% of the entire payload mass (without the thermal components). The TRL and flight heritage are not listed for mass components with an “Estimate” mass maturity since these masses are based on percentages of the total mass. However, the Science Payload thermal components and secondary structures will be constructed using the materials, coatings, and processes that have been used on past NRL programs such as SECCHI and VAULT. Table J-1. Basis of Estimate Definitions Maturity Contingency Code Example Basis (%) Estimate E 20.0% New Development Hardware Calculated C 15.0% Determined by Calculation or Analysis Modified MD 15.0% Modified Flight Hardware Detailed D 10.0% Detailed Design or Fab Drawings Complete Design Heritage H 5.0% Existing, Build to Print Flight Hardware Specification S 3.0% Off-the-shelf Flight Proven Hardware from Vendor (Off the Shelf) Measured MS 0.0% Measured Mass/Power of Flight or Qualification Model
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Table J-2. SoloHI Master Equipment List
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