SDO Systems Retreat - Stanford University

SDO Systems Retreat - Stanford University

Solar Dynamics Observatory (SDO) Spacecraft Design and Operations Overview Mission PDR David Ward SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 1 Agenda Spacecraft requirements overview/driving requirements status Spacecraft trades and major changes since SCR Overview of architecture and spacecraft subsystems Observatory operations concept Technical resources Mass, power, pointing, jitter, data capture, data completeness, propellant Issues Conclusion SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 2 Status of Spacecraft Requirements Top-level spacecraft requirements (defined in MRD) have remained relatively stable since SRR/SCR. Among the changes, many stem from the replacement of SHARPP with AIA, as well as lessons learned from the SRR/SCR. Removal of KCOR eased absolute pointing and particulate contamination requirements Removal of the OFS from EVE eliminated local magnetic field requirements Jitter requirements have been clarified through detailed discussions with the instrument teams SHARPP data rates have been reallocated, allowing EVE to eliminate compression Technical resource allocations were modified to share some project margin with subsystems Since Autumn, more emphasis has been on Level 3 (and 4) requirements, further detailing subsystem requirement allocations and design decisions Subsystem requirements began draft development in the Summer/Fall, and the project has been baselining documents after subsystem PDRs to allow for peer review of requirements Subsystem requirements include verification matrices that follow the MRD style Requirement, traceability, rationale, assignment and verification method are tracked for all requirements At Level 4 (component specs/SOWs), documents are being created for standalone use, so that vendors

do not have to search through multiple document to understand all of their requirements Component level specs are reaching their final draft stages (one is baselined), and are nearly ready for this Springs procurement activities At this point, spacecraft and subsystem requirements are in good shape to proceed into detailed design SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 3 Mission Design Drivers The primary mission design drivers have remained constant through preliminary design: High data volume, coupled with tight requirements on data loss and degradation Drives requirement for high speed science data bus and Ka transmitter to downlink data Also derives requirement for dedicated SDO ground station and high gain antenna control Significant effort dedicated to Ka Transmitter breadboarding, data loss analysis and budgeting since SCR Additional effort placed on HGAS calibration on-orbit, in order to improve RF gain, and thus bit error rate Geosynchronous orbit Drives launch vehicle and propulsion system requirements, and places SDO in high radiation environment and electrostatic return environment for contamination Propulsion design has been modified to provide a backup GEO injection approach, and has converged on a more traditional GEO injection design Detailed radiation analysis of preliminary parts list is underway, as well as design mitigation as necessary Electrical systems will place a continued emphasis on grounding, on-orbit ESD, Common Mode Noise in recognition of special challenges of the orbit Long mission life (5 year requirement, 10 year goal) Drives reliability (especially of mechanisms), redundancy, & radiation requirements Preliminary designs use mechanisms (gimbals, reaction wheels, filter wheels, etc) that have proven through life test and flight data their ability to meet the life requirements Tight pointing, jitter, and coalignment requirements Driving jitter requirements clarified since AIA brought on-board Trade to add a Guide Telescope for each AIA Science Telescope minimizes differential flexibility risk and brings jitter approach in-line with proven TRACE approach Preliminary analysis of jitter and pointing budgets shows requirements can be met using baselined design SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 4 Design Trades Performed Since SCR Trade Considered Orbit design Propulsion module

design Propulsion design Propulsion Module thermal design Solar array sizing Omni antenna location Ka transmitter design: integrated vs componentized Reaction wheel make/buy Number and location of AIA guide telescopes EVE science data bus allocation AIA science data bus allocation HMI science data bus allocation Selected Option Rationale minimizes momentum buildup during GTO phase, allowing for use of available Reaction Wheels separation perigee at 300 km nominal insertion at 185 km without requiring thruster control during perigee passes stationkeeping can be performed with thrusters only in one direction, allowed for Propulsion thrusters all anti-sunward; two thrusters fire in three different Module to completely be a separate unit from spacecraft module, for parallel I&T; two tank stack tank stack directions, four tank module eased development of PM, moved slosh mass onto centerline once all thrusters anti-sunward, contamination concerns from MMH significantly reduced; less Dual-mode N2H4/NTO Traditional MMH/NTO biprop complex design; higher Isp of ACS htrusters allows for a backup to orbit in event main engine biprop/monoprop fails Hybrid of "toasty cavity" traditional design with around tanks and internal and simplifies heater design around tanks, reduces risk of a problem with an unaccessible individual wrapping; only a individually wrapped thruster heater/thermostat, uses an acceptable amount of heater power compared to toasty cavity alone toast cavity lines 5.8 m^2 and PSE small savings in solar array size did not outweigh cost/complexity/risk of abandoning MAP 7.7 m^2 modifications heritage DET design Omni boresights angled X axis boresights have antenna pattern nulls located such that stationkeeping burns would be either +/- X or +/- Z boresights between X and Z in XZ plane affected, Z axis boresights would place nulls during eclipse Integrated design Options Considered Componentized (components baseline design is integrated transmitter, for best noise performance and ability to trade-off connected via coax) performance among various components as more is learned during design Out-of-house design Four AIA Guide Telescopes. Each assigned to one Science Telescope In-house design At least two different commercially available Reaction Wheels meet the requirements of the mission, including mass balancing; "buy" approach can select a design with existing life-test, rather than requiring a life-test for a new design Two AIA Guide Telescopes, eliminates concern about local flexibility between GuideTels and SciTels, makes jitter attenuation shared among the four AIA design very similar to TRACE Science Telescopes 7 Mbps 2 Mbps eliminates compression requirement for EVE, simplifying electronic design 67 Mbps 58 Mbps minimizes lossy compression for AIA, improving science quality

55 Mbps unchanged SDO Preliminary Design Review (PDR) March 9-12, 2004 HMI already using lossless compression, no change required Spacecraft Overview Page 5 Major Changes Since SCR Replacement of SHARPP instrument suite with AIA instrument At a concept design level, many of the spacecraft interfaces (power services, high speed bus, 1553) moved directly from SHARPP allocations to AIA; additional work ongoing to flesh out detailed interfaces (mechanical/thermal/electrical/data/pointing) Minor redesign of the Instrument Module (IM) performed to give AIA telescopes more integration room; IM now more of a square than offset rectangle as was required by side-mounted SPECTRE Decision to mount four Guide Telescopes directly to AIA Science Telescopes eases structural flexibility concern from separate mounting Propulsion system trades revisited, resulting in change to more traditional MMH/MON-3 biprop design Separation of Solar Array and High Gain Antenna deployment functions Deployables team traded several different actuators for these functions; baseline design uses QUIKNUT actuators for both the Solar Arrays and High Gain Antennae Each array and HGA boom is on a separate deployment circuit; S/As autonomously deployed at separation, HGAs deployed individually by ground command after Observatory is power-positive Increased Solar Array size without affecting HGA coverage GN&C team (ACS, Propulsion and Flight Dynamics) revisited original configuration that required thrusters on -X and either Y or Z axes, instead moving all thrusters to -X direction Resulting configuration significantly eased contamination concern and allowed consideration of more traditional MMH/MON-3 system In addition to programmatic benefits, higher Isp ACS thrusters gave a backup to the main engine Array area increased from 5.8 sq. meters to 7.7 sq. meters to account for reduced array output at lowest end of bus voltage range HGA six month coverage protected without additional boom length by optimizing taper shape of arrays to account for distance between HGA and arrays Solar arrays oversized compared to the load to minimize possibility of future array redesign and resulting impacts Elimination of EVEs OFS and EVE Electronics Box (EEB) move onto IM Removal of OFS eliminated magnetic field sensitivity/requirements, leaving only dipole as a mag requirement EEB is easily accommodated by larger IM (resized to accommodate AIA); move saves on intra-instrument harnessing, but places more difficult radiation burden on EEB SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 6 Design Changes: SHARPP/AIA Some minor spacecraft redesign resulted from replacement of SHARPP by AIA Removal of side-mounted SPECTRE allowed for IM to be squared off, providing more room for AIA telescopes

AIA telescope overhang acceptable; does not interfere with HMI, EVE FOVs (more difficult if either had KCORs wide stray light FOV) AIA takes responsibility for controlling Guide Telescopes; pointing budgets very similar AIA planning to use same Camera Electronics/CCD as SHARPP; thermal radiator requirements very similar High Speed Bus for science data (shown below) is a simple interface replacement, and some reallocation of SHARPP data rate between AIA and EVE With KCOR removal, particulate contamination requirements eased 1 2 4 Ports 3 2 Active 4 AIA 1 4 2 Primary 5 Ka Comm 3 6 1 2 4 Ports 3 2 Active 4 SDO Preliminary Design Review (PDR) March 9-12, 2004 KCOR AIA PDR 1 2 4 Ports 3 2 Active 4 EVE Magritte SPECTRE Primary Ka XMTR HMI 1 4 2 Redundant 5 Ka Comm 3 6 POST SCR Config. 8/03 Config. 12/03 Redundant Ka XMTR Spacecraft Overview Page 7 Design Changes: New Thruster Locations SDO Preliminary Design Review (PDR) March 9-12, 2004 Four pairs of canted thrusters surround the main engine, with each pair assigned to separate isolation banks The canting allows diagonal pairs to be used for X control, in addition to adjacent thrusters being used for Y and Z control

Since the eight ACS thrusters are also biprop engines, they can be used to get to GEO in the event the main engine fails Thrust direction of all nine thrusters away from instruments, easing the contamination risk presented by propellant Spacecraft Overview Page 8 Observatory Mechanical Configuration Composite Instrument module minimizes thermal distortion, places instrument radiators outboard and with clear thermal field of view HMI AIA AIA (which uses GT signals for IMC) all on one face, HMI, EVE (which do not) on the other EVE AIA EVE HMI Two short tapered arrays, cell side out when stowed Spacecraft bus module provides Faraday cage and radiation shielding for s/c and instrument components Internal propulsion module, allows for parallel integration and test flow early Redundant High Gain Antennae (HGA) at the end of rigid booms. Each antenna can be used continuously for ~ 6 months/year (scheduled antenna handovers twice/year) SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 9 SDO Propulsion Module Fuel Tank Pressurant Tanks Oxidizer Tank (Inside Cylinder And Frustum) Fuel/Oxidizer/ Pressurant Control Modules Thrusters (4 Places) Ka Transmitters and Switch (Panels Removed) SDO Preliminary Design Review (PDR) March 9-12, 2004 Fill and Drain Valves Main Engine Spacecraft Overview Page 10 SDO Spacecraft Subsystem Overview Command & Data Handling Attitude Control Jitter performance at focal plane to <0.5 (3)), calibrated pointing accuracy of 10 (3)) via zero-momentum, three-axis

control with Reaction Wheels Hardware command decoding for computer-free recovery Provides continuous 130 Mbps high-speed interface between Instruments and Ka-band RF system Star Tracker, Inertial Reference Unit, and Guide Telescope used for target/attitude determination Momentum unloading monthly with thrusters Provides Spacecraft Processor for high-end Attitude Control algorithms, command/telemetry processing Up/Down card provides interface to S-band RF system Communications Ka-band transmitter through two High Gain Antennae to downlink science data S-band Transponders connected to Omni antennae for receipt of ground commands (2 Kbps) and telemetry downlink (64 Kbps) via SDO Ground Station, USN, TDRS Supports orbit determination via turnaround ranging Power Two Solar Arrays for string-fault-tolerant power generation supporting a 1450 W load One Lithium Ion battery (100 A-hr) for launch and 72 minute eclipse survival at nominal load Power switching is distributed, with high current switches in PSE and low current distributed to various subsystems Software Complex algorithms computed on central processor, including ACS, Stored Commanding, Solid State Recorder, and Fault Detection & Correction ACS, C&DH, HGAS and Power each have smaller embedded processors for power switching, housekeeping telemetry generation, and subsystem-specific applications (Safehold, Load-shedding, Thermal Control) Common software used for RTOS, 1553 RT, Time, Memory Load/Dump, Power Switching, etc on all SDNs SDO Preliminary Design Review (PDR) March 9-12, 2004 Propulsion MMH/MON-3 bipropellant design to raise orbit from GTO, perform E-W S/K, unload momentum 445N (100#) engine used for GTO (with 22N (5#) ACS thruster backup) All thrusters on aft end of Observatory to limit contamination, improve observatory modularity Mechanical & Mechanisms Designed for EELV (Delta IV 4040 or Atlas V 401)

Octagon structure with electronics mounted to inside of exterior walls for better thermal heat rejection On-orbit symmetry to minimize momentum buildup Deployable solar arrays and high gain antennae with uninterrupted coverage on one antenna for 6 months/year (no handovers needed) HGA pointing to 0.25 to support Ka link margin Continuous antenna pointing on same HGA (slip rings) Thermal Combination of passive/active design Software controlled operational heaters (optical bench) Thermostatic control of survival heaters Hybrid approach or toasty cavity and individual line heaters to minimize risk in propulsion thermal design Spacecraft Overview Page 11 SDO Electrical Architecture 28V Power to ACS sensors, actuators &heaters HMI ACE A Power Switching CSS DC-DC Converter HMI Optics & CEB RT HMI Inst Electronics RT BC ACE SDN AIA RT Engine Valve Driver boards MEGS RT Propulsion ST #1 RT ST #2 RT to Up/Down B Synchronous Serial Bus ESP DC-DC Converter H/W decoded cmds EVE IEM (incl. SDN) Pwr Switching to High Gain

Antennae Thermistors, HGA sensors 28V power GCE CDH A RT IRU Waveguide Switch from Ka Comm B to S- Band 4 Optics & CEB AEB Bulk Memory & DC/DC Converter Uplink/Downlink S Band SDN EVE Prop Pyro board Ka XMTR B DC/DC Converter S/C Processor RT GTs RWA I/O Ka XMTR A High Speed Data Ka Band Pwr Switching DC-DC Converter 1553 Bus Gimbal Interface RT RW #1 from Instruments RW #2RW #3 RW #4 DC-DC Converter RT DC-DC Converter RT DC/DC Converter RT ACE B 28V Power to ACS sensors, actuators & heaters 28V power 28V power Solar Array Solar

Array Battery SDO Preliminary Design Review (PDR) March 9-12, 2004 High Speed Data Ka Band DC-DC Converter to Up/Down A Pwr Switching Bulk Memory & DC/DC Converter S/C Processor Power Switching Gimbal Interface CDH B BC PSE Housekeeping SDN RT PSE SDN 3 Output Modules Battery Module Solar Array Module ACE SDN Deploy Circuits CSS Solar Array Module RWA I/O 3 Output Modules Prop Pyro board PSE SDN DC-DC Converter Engine Valve Driver boards to Ka- Band RT RT Housekeeping SDN Synchronous Serial Bus 28V Power to Gimbal drives, Instrument Module thermal control Uplink/Downlink S Band SDN H/W decoded cmds DC-DC Converter S XPNDR A Pwr Switching S XPNDR B

28V power to SBC, Ka Comm, S XMTR from S Comm A 3 dB Hybrid to Omnis Spacecraft Overview Page 12 Common Design Elements Subsystem Data Node Subsystem Power Node Serves as the embedded processor for many of the spacecraft avionics boxes Since SCR, split into two boards to provide all of the common power requirements Uses a Motorola RH-CF5208 ColdFire processor for processing Provides MIL-STD-1553 for communications with the spacecraft processor and cPCI for backplane communications The Power Conversion Card (PCC) provides DC/DC converters for 2.5, 3.3, 5, and 15 V, and provides a staggered enabling for those voltages to deal with FPGA power-on issues The PCC contains voltage monitoring circuitry to provide a power-on reset signal to the other electronic cards in the event one of the regulated voltages exceeds limits The PCC provides an external interface to command a power-on reset for the S Comm Card and the PSE, which are unswitched The Low Power Switch Card (LPSC) provides 16 switched services (8 @ 1A, 8 @ 2A) for further distribution of 28V power In the event of a converter regulation anomaly, the switches are configured to hold state as long as the 28V power is still supplied (all LPSCs are on switched 28V services) Provides external interface to command a processor and backplane reset without changing the status of the other circuitry Also provides passive and active analog conversion circuits Common SDN software includes RTEMS RTOS and GSFC-developed software bus Prototype unit completed in Fall 2003, allowing for ringout of HW/SW interfaces prior to subsystem breadboard deliveries

SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 13 Additional Spacecraft Design Highlights Preliminary design progression results in detailed allocation of requirements Observatory-wide interfaces like power switches and 1553 bandwidth allocation have preliminary designs, which show adequate spare services for PDR Orphan functions, such as heater/thermistor services, the waveguide RF switch driver, deployment pots and separation switches have all been assigned to avionics Propulsion redesign added pyro valves, whose drivers have been assigned to the ACE (as were the thruster and isolation valve drivers) Twelve of sixteen hardware decoded commands assigned; they include processor and full resets for the S Comm cards and the PSE sides, spacecraft processor resets and a command to switch either spacecraft processor from a non-BC mode (TBD, either RT or Standby) to BC Longer development items such as the SDN core and components of the Ka Transmitter have breadboard/prototype designs to work out design details Integrated Ka Modulator breadboard completed in the Fall, followed by Ka Solid State Power Amplifier breadboard completed this Winter and preparing for performance and life degradation testing Reaction Wheel and Inertial Reference Unit interfaces simplified (eliminate 1553) RWAs are required for Safehold, and IRU may be when detailed design is complete Since 1553 will not be used as a Safehold data interface, decision was made to simplify the component designs by only using one data interface SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 14 Observatory Operations Concept Allows Functional and Operations overview of system design as a part of the Big Picture of requirements, Implementation Approach, and Ops Concept rather than as isolated requirements Mission operations are arranged into five time-sequenced phases, which include detailed modes or activities that are also described in the following charts Launch and Acquisition Phase In-orbit Checkout & Orbit Circularization Phase Instrument Commissioning Phase Science Mission Phase Normal Mode Periodic Calibrations/Housekeeping

Eclipse Mode Stationkeeping & Momentum Management Safehold & Emergency Modes Disposal Phase The activities listed under Science Mission Phase are described in the context of that phase, but the capabilities are not limited to only that phase The safeguard capabilities described in Safehold and Emergency Modes exist in every phase Momentum management will be performed in every phase, but the once per four week constraint does not apply to the first two phases Eclipse mode preparations are similar for all phases; for the early phases some components not yet powered SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 15 Launch and Acquisition Phase covering pre-launch configuration until Observatory is power-positive and pointing at the sun Observatory kept in low-power configuration: Instruments (& decontamination heaters) off, redundant units off, science data components (Ka Comm, Ka XMTR and Star Trackers) off Launch until separation is approximately 45 minutes to a separation altitude of 300km Separation altitude increased to reduce high momentum buildup due to atmospheric drag (at 185km) Transmitter powered minutes before separation to allow for telemetry at separation (now expected through existing ground network station at Overburg, South Africa, Perth or Dongara Australia) TDRSS available as contingency/backup Autonomous Solar Array deployment and Reaction Wheel power application at separation, based on separation signals backed by software sequencer Attitude Control System acquires sun in 45 minutes from separation rates to within 15 of the sunline Rate damping can begin immediately after separation, but CSS sun errors are ignored until array deployment is sensed Following discussions with KSC, SDO LV IRD specifies [0.25,0.25,0.25]/s, which will eliminate need for thruster-based momentum unloading until after Observatory is power-positive (capability for ground commanded unloading still exists in the event of a separation anomaly) Once the observatory is power-positive: Instrument CCD decontamination heaters powered on (Instruments remain off) Power on GCE (includes Housekeeping card) to provide additional thermistor data Deploy the HGAs nominally within 2 hours of separation

SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 16 Launch and First Orbit Timeline SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 17 In-Orbit Checkout Phase used during first weeks to checkout and calibrate Observatory Phase is concurrent with orbit circularization phase Spacecraft components brought on-line, and capabilities/modes checked Hot backup ACE powered on High Gains deployed within hours after separation ACS/Propulsion Checkout and calibration o Safehold checkout o Sensor Checkout/Calibration IRUs, Star Trackers o Inertial Hold / Slew Capability checkout prior to fist planned maneuver o Thruster checkout prior to first planned maneuver (phasing for 5lb thrusters) Observatory communications via external ground networks and SDO ground station SDO dedicated ground station not available for continuous coverage until 3 rd apogee maneuver Instruments not powered on until all large apogee maneuvers complete Maintains power margin in the event of an anomaly Instrument CCD decontamination heaters remain on Instrument doors remain closed. High rate science data system (Ka-Comm, Ka-XMTR, HGAs) brought online for HGA pointing calibration and system checkout once at GEO slot SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 18 Orbit Circularization Phase used during first weeks to circularize the orbit from the GTO and place SDO in its geosynchronous slot at 102W. Phase is concurrent with In Orbit Checkout phase Four (4) large Apogee Motor Firing (AMF) and three (3) small Trim Motor Firing (TMF) maneuvers are planned to place SDO in its final geosynchronous slot AMF maneuvers use 445N (100#) thruster, TMF maneuvers use 22N (5#) thrusters Approximately 2 weeks to complete Total maximum duration for any maneuver activity will be less than ~90 minutes

Maximum slew time of 20 minutes before/after, settling, 50 minute maximum Delta V Observatory may be pointed to any orientation during maneuver, so power, thermal, other designs must take 90 minute off-pointing as a requirement Maneuvers are not time critical If a maneuver is aborted or missed it can be made up later with no penalty Observatory communications via external ground networks and SDO ground station Thruster burns must be started and completed within view of one (or more) ground station Consideration given to slight delay of apogee burns until Observatory in sight of station Commands for maneuvers uploaded to Absolute Time Sequence buffer, rather than singularly commanded from ground SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 19 In-Orbit Checkout & Orbit Circularization SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 20 In-Orbit Checkout & Orbit Circularization SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 21 Instrument Commissioning Instrument calibration and commissioning begins once on-station and lasts 30 to 60 days Instruments are powered on (if not already) and optics doors are opened Observatory communications through SDO ground station for S and Ka band High rate science system brought on line (Ka-Comm, Ka-XMTR, HGAs) HGA calibration is performed to remove static misalignments SDO Ground Station tracks RF power while HGA performs raster slews Instruments begin producing science data Science data distributed directly from SDO ground station to SOCs Instrument operations support from both MOC and SOCs Spacecraft supports instrument calibration roll maneuvers and off point maneuvers Maneuvers similar to periodic instrument calibration maneuvers described later SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 22

Nominal Mission Mode (Science Phase) Expected to be phase that mission stays in 99% of time once at GEO, with few operational activities/interruptions normally planned Ka-band science data is downlinked through SDO ground station and distributed to SOCs on continuous basis S-band housekeeping data is collected by ground site and distributed to MOC, which further distributes data to SOCs Nominal downlink rate is 64 kbps to the SDO ground station (data on RF carrier) Twice daily periods of s-band omni antenna RF interference may degrade H/K data S-Band data rate may be reduced to improve link margin during interference times Orbit tracking operations performed two consecutive days a week 6 passes (30mins each) each day from the SDO ground station and 1 pass (30 mins) each day from an external ground station (Hawaii) potential to reduce tracking to bi-weekly RF reconfiguration required for tracking, data placed on subcarrier and data rate lowered Instruments SOCs will have a normal window each weekday to command Instruments and uplink loads with all commands passing through MOC to ground site Anticipate weekly loads since instruments are full sun viewing with routine operations Contingency command periods if on-duty FOT member(s) are contacted and bring up command link Spacecraft data recorder maintains circular buffer with 24 hours of housekeeping data in order to capture anomalies in case of data loss Attitude control system autonomously points reference boresight to sun and maintains proper rotation about sunline Proper orientation is achieved by Inertial slew to sunline using Star Tracker attitude, then switching to Guide Telescope for science pointing SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 23 Nominal Mission Phase (Science Phase) SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 24 Periodic Calibrations/Housekeeping There are several periodic interruptions to the nominal science mission mode: Stationkeeping and Momentum Unload maneuvers, Instrument Calibration (Roll and Offpoint) maneuvers, eclipses (earth and moon), HGA handovers All scheduled interruptions which cause science data loss are included in the Data Capture Budget Twice a year HGA Handovers

Each HGA has an unobstructed field of view for approximately six months Need to turn on Ka-Transmitter a few hours before handover to stabilize TCXO Periodic HGA calibration may be required to maintain HGA pointing requirement Thermal effects (to HGA boom) may degrade HGA pointing RF signal strength degrades rapidly as HGA boresight pointing drifts outside the nominal antenna beamwidth Instrument teams have identified periodic calibration activities Roll maneuvers to observe solar shape Off-point maneuvers for flat fielding and optical distortion calibration Alignment adjustments to align instrument to reference Most instrument calibrations are infrequent, but AIA Guide Telescope/Science Reference adjustments, coordinated with HMI Alignment Leg adjustments, planned for up to once every two weeks SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 25 Instrument Calibration Maneuver Details HMI: Off point twice a year: up to +/- 30 arcminutes (~ solar diameter) about twice a year: ~20 positions/5 minute dwell at each position. This is a scan and step pattern. 360 degree roll twice/year: 16 positions/22.5 degree steps/15 minute dwell at each position. Alignment adjusts anticipated up to every two weeks: adjust HMI mounting legs to keep instrument aligned with reference (Guide Telescope) and keep Image Stabilization System (ISS) in range EVE: Cruciform off-point scans quarterly: 180 arcmin mapped at 3arcmin per step with dwell at each position of 30 seconds (total of 60 dwell points) FOV Maps quarterly: 25 point, 5x5 map, 5arcmin/step covering +/- 10 arcmin each axis, hold each position for 60 secs then advance AIA: (preliminary, somewhat based on similar SHARPP conversations) Roll & off point calibrations perhaps twice a year (match to HMI timeline) GT calibration (frequency is TBD), scans of few arcminutes wide, one in pitch and one in the yaw direction. ACS will use a Star Tracker for knowledge, so we can complete these maneuvers without the GT signals. High-rate science data is needed during the dwell points in the calibration maneuvers High-rate science data is not needed during the calibration slews, only during dwell periods: this applies to both off-points and rolls. Must ensure HGA coverage while attitude changes for slews requires coverage planning Maneuver sequences require spacecraft and instrument coordination FOT will direct activities and produce coordinated activity plans and time-tagged command loads Maneuver sequences will be performed by on-board time-tagged command loads SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 26 Eclipse Observatory requirement in this phase is to survive and minimize impact on science operations Some Observatory configuration changes are made at the start of eclipse Instruments are left powered, and instrument and spacecraft (optical bench) thermal/heater power increases to minimize thermal distortions during eclipse Attitude control reverts to Star Trackers, due to loss of Guide Telescope signal

Arrays sized to fully recharge battery before next eclipse Prior to the start of eclipse season the battery charge control algorithm will be set to increase battery state-of-charge to 100% pre-eclipse At the end of the season the battery will be returned to a reduced state-of-charge to prevent overcharge Observatory Failure Detection and Correction will be configured for eclipse Deeper than normal battery discharge Safehold designed to respond properly without coarse sun error input Ka-Band subsystem will continue to perform Ka communications may be degraded due to thermal effects on antenna booms Instruments will continue to produce science data and expect to receive it on a best effort basis Recovery from eclipse state budgeted at 1 hour after eclipse exit (for HMI science data) Recovery defined as achievement of pointing/alignment, thermal requirements after eclipse SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 27 Eclipse Timeline SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 28 Stationkeeping/Momentum Management Required operations to keep SDO within its orbit slot at 102 W and to maintain Observatory angular momentum near zero To meet data capture requirements, this phase is only budgeted to interrupt science once/month Stationkeeping burns (Delta-V) alone require a twice yearly interruption Momentum management (Delta-H) will occur approximately monthly (actually every 4 weeks) Requirement on spacecraft to handle 5 weeks period between momentum unloads, which allocates 4 weeks for nominal operations and 1 week for Safehold operations One hour allocated for science interruption due to stationkeeping and 30 minutes for momentum management During Delta-H spacecraft remains sun pointing but pointing control is +/- 5 During Delta-V spacecraft may be off pointed up to 45 from the sun-line for up to 30 minutes Major reconfiguration (instrument power, Ka RF) not necessarily warranted unless power constraints require non-essential power to be reduced Since thrusters moved to bottom deck, instrument doors do not need to be closed during maneuvers (to avoid contamination effects) Given nature of operation, the SDO (or alternate) ground site is maneuver critical. Maximum offset of 15 from the XY plane to minimize sun on instrument CCD radiators

All stationkeeping and momentum management burns qualify as critical operations that must be viewed by the ground. Can roll the spacecraft or delay burn time to ensure good communications (null avoidance, improve omni antenna coverage) for SK (or momentum dump) maneuvers. An alternate ground station may be substituted for the SDO site if the SDO ground site is unavailable Instrument teams will receive a 1553 warning message prior to and upon completion of each maneuver Instruments will take pre-described action upon receipt of critical event notification commands SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 29 Typical Stationkeeping Timeline SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 30 Safehold/Emergency Modes Several capabilities will exist on the Observatory for safing in the event of an anomaly Critical event notification commands have been identified to inform instruments of current or pending conditions. Fault Protection software (FDC/TSM/RTS) in main spacecraft processor to respond to anomalous housekeeping telemetry For attitude control anomalies, Observatory will drop into a simpler sun-pointing control mode either controlled by spacecraft processor (Sun Acq) or by independent ACE SDN Independent ACE safehold can be commanded by ground or by spacecraft processor as response to FDC actions Loss of Im OK communications between ACE and s/c processor will cause ACE to enter Safehold Safehold works without ground intervention until momentum limits reached o Momentum capabilities sized for one week of control before ground commanded unload performed o Safehold will not autonomously fire thrusters to unload, avoiding possibility of tumbling spacecraft o Safehold and Sun Acq are sun pointing. Same orientation as nominal science pointing o Some consideration being given to a Safehold command to effect a coarse roll about sunline, in order to move out of communication null For power anomalies, spacecraft processor and PSE have layered load-shedding algorithms to reach lowest power state Spacecraft processor can power-off individual components switched by various LPSCs (Star Trackers, Ka Comm, optical bench thermal control) Independent PSE load-shedding powers off services at a lower state of charge/battery voltage (can only open switches at Output Module level, turning off instruments, possibly the GCE, Ka Transmitter, if not already powered off by spacecraft processor) Commands sent from main processor across the 1553 bus Commands indicate safehold entry, pending load shed power off, eclipse entry, etc Loss of the time distribution message across the 1553 can be used to indicate loss of main processor or 1553 Uplink communications path is redundant and receiver and uplink card are on unswitched power Hardware commands decoded in the uplink card hardware allow critical subsystem reconfiguration to recover nominal on-board communications SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 31 Disposal At end of mission, NASA policy requires disposal of SDO into an orbit that wont interfere with other spacecraft Increase altitude to >300 km above GEO orbit The actual de-orbit altitude is GEO + 300 km + X, where X is a function of the spacecraft mass and crosssectional area. Operations similar to orbit circularization at beginning of life

In order to ensure enough power for operations, instruments and science-oriented spacecraft components will be powered off PDR Orbit Debris Assessment has been completed by Josephine San, and is in SDO CM review SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 32 Technical Resources Management At project level, the following technical resources are being managed: Mass, power, alignment/pointing, propellant, data capture, science data bus data rate, bit error rate (to meet data completeness), 1553 bandwidth, RF link margins Mass and Nominal Power allocations baselined prior to SCR, with allocation increases made in late October, along with baselining Eclipse Power Original allocations matched SCR estimates, with process requiring CCRs for any increases in allocation as a way to slow resource growth By October, resource growth had slowed considerably, allowing for project to allocate some of its reserve to each subsystem and instrument team to be held at their level (per a SCR recommendation) Project still holds reserve for each configured budget to maximize likelihood of hitting percentage targets (25% margin at PDR, 15% margin at CDR) Four additional power states will be baselined: Launch, Orbit Raising, Survival, Stationkeeping As shown on the following charts, SDOs technical resources show an appropriate level of margin for PDR design maturity Mass and power well above 25% margin for all modes, and are measured against worst-case failure conditions (12% decreased the solar array capability in the normal mode case and 13% increased current draw in the eclipse case) Pointing/alignment/jitter budgets are challenging, but achievable using available components/methods Data capture/completeness well analyzed, and science/housekeeping bus bandwidths correctly allocated Propellant still has margin with a worst-case stackup, and an option exists to build more margin if necessary In the following charts, the terms Project reserve and margin are not interchangeable Project reserve = total capacity - total allocations: the allocation margin held at the project level Margin = total capacity - current best estimates: the actual measure of project resource margin Values listed on the next page are margin SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 33 Technical Resources Status/Margins SCR PDR CDR Flight CBE

Margin Dry mass 30% 25% 15% 0% 29.4% Project holds 20% reserve, rest allocated to instruments and subsystems Power Normal Mode 30% 25% 15% 0% at EOL 30.0% sunlit mode, all instruments on, assumes lower voltage due to one failed cell margin against 60% DoD, assumes one failed batttery cell, 72 minutes, normal power configuration margin against 60% DoD, assumes one failed batttery cell, assumes 120 for entire phase, 45 minutes launch to separation margin against 60% DoD, assumes one failed batttery cell, 90 minutes entire phase, 50 minute burn margin against 60% DoD, assumes one failed batttery cell, 45 degrees off sunline margin against 80% DoD, assumes fault at exit of eclsipe, followed by 30 minute recovery (also one failed batttery cell) steady-state power generation margin against Survival Mode power Resource Eclipse Mode 27.0% Launch Mode 28.3% Orbit Injection Mode 25.3% Stationkeeping Mode 139.9% Survival Mode 34.3% Survival Mode 64.2% Comments ACS thruster backup 5.0% Main engine, w-c stack ACS b/u, RSS stack 3.8% 1.5% worst-case stackup includes -3% Isp, maximum ACS control and momentum buildup, worst case mixture ratio, disposal propellant propellant margin at 3200 kg, nominal performance values propellant margin at 3200 kg, nominal performance values, except Isp for ACS thrusters during Orbit Injection all penalties added in worst-case stackup knowledge uncertainty, mass ratio and launch dispersions RSS'd 60% 75% worst case estimate is ACE SDN 65% 58% based on specified SBC MIPS performance, thought achievable in industry

assumes 12 MIPS processing, which may require oscillator change from prototype Propellant positive margin with 3 sigma usage stackup Main engine 10.1% Memory Spacecraft Processor SDN's 50% 50% CPU Throughput Spacecraft Processor SDN's 50% 1553 Bus Bandwidth 20% 15% 10% 5% RF Link Margin Ka Downlink S Uplink S Downlink on Carrier 3 dB 3 dB 3 dB 3 dB 50% S Downlink on Subcarrier 40% 40% 25% 25% 18% 4.2 dB 16.5 dB 4.3 dB 3.3 dB assumes 64 kbps downlink to SDO Ground Station assumes 32 kbps downlink to SDO or commercial Ground Station Detailed breakdowns of mass, power, 1553 budgets available in backup charts. SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 34 Mass Budget Details Project reserve is healthy for additional allocations if necessary Subsystem CURRENT ALLOC CURRENT BEST ESTIMATE Mechanical 575.00

542.34 Pow er 103.00 96.50 ACS 103.00 97.30 Propulsion (Dry Mass) 138.00 131.83 C&DH 28.00 26.10 S Comm 9.60 7.40 Ka Comm 22.00 16.75 Mechanisms 50.00 45.60 Thermal 49.00 45.30 Electrical 154.00 137.00 HMI 71.00 64.30 AIA 130.00 120.19 EVE 52.00 48.44 Total Dry Mass 1484.60 1379.05 Propellant and Pressurant 1415.00 1415.00 Total Mass Allocation/Estimate Mass Capability to GTO Dry Mass Margin vs Allocation/Estimate Dry Mass Margin 2899.60 3200.00 300.40

20.23% 2794.05 SDO Preliminary Design Review (PDR) March 9-12, 2004 405.95 29.44% Total Mass vs Capability Mass (kg) % CBE Margin 29.44% 3500.00 3400.00 3300.00 3200.00 3100.00 3000.00 2900.00 2800.00 2700.00 2600.00 2500.00 Jan03 Mar03 May03 Jun03 Aug03 Oct03 Nov03 Jan04 Mar04 Date Dry Mass Margin 40.00% Percentage Margin Mass budget shows comfortable margin Total Mass (kg) 2899.6 35.00% 30.00% 25.00% 20.00% 15.00% Dec-02 Mar-03 Jun-03 Oct-03 Jan-04 Apr-04 Date Budget based on 3200 kg separation mass. Margin measured against dry mass, given a known propellant load Spacecraft Overview Page 35 Sunlit Power Budget Details Subsystem Current Allocation Normal Mode (W) CBE Normal Power (W)

Power 54.00 52.0 Electrical/Harness 22.50 21.1 ACS 251.00 241.6 Propulsion 11.00 10.0 C&DH 95.00 95.0 S Comm 48.00 45.8 Ka Comm 62.00 53.0 Mechanisms 41.00 40.5 Thermal 155.00 148.0 HMI 111.00 92.3 AIA 135.00 111.2 45.00% EVE 76.00 67.5 40.00% LOAD BUS TOTAL 1061.50 978.0 Efficiency Losses 159.23 143.4 TOTAL POWER 1220.73 1121.3

1450.00 24.78% 1653.5 47.46% 1458.0 30.02% Nominal Mode Power vs Minimum Capability 1500.00 1400.00 1300.00 1200.00 Watts DKW 1100.00 1000.00 900.00 800.00 Jan-03 Mar-03 May-03 Jun-03 Aug-03 Oct-03 Nov-03 Jan-04 Mar-04 Date Project reserve is healthy for additional allocations if necessary Nominal Power Generation Nominal Power Generation Margin One-Fault Power Generation One-Fault Power Generation Margin SDO Preliminary Design Review (PDR) March 9-12, 2004 50.00% Percentage Margin Normal mode shows comfortable margin Nominal Mode Power Margin 35.00% 30.00% 25.00% 20.00% 15.00% Jan-03 Mar-03 May-03 Jun-03 Aug-03 Oct-03 Nov-03 Jan-04 Mar-04 Date Budget based current solar array size, with generation capacity limited by bus voltage, assuming one failed cell (worse failure than failed string or failed PWM.) Spacecraft Overview Page 36 Eclipse Power Budget Details CBE Eclipse Power (W) Power 54.00 52.0 Electrical/Harness 21.50 19.3 ACS 251.00 241.6 Propulsion 11.00 10.0 C&DH

95.00 95.0 S Comm 48.00 45.8 Project reserve is healthy for additional allocations if necessary 65.00 60.00 55.00 50.00 45.00 40.00 Oct-03 Ka Comm 62.00 53.0 Mechanisms 41.00 40.5 Thermal 200.00 148.0 HMI 122.00 102.3 40.00% AIA 176.00 147.2 35.00% EVE 76.00 68.5 LOAD BUS TOTAL 1157.50 1023.2 Efficiency Losses 14.00 11.8 TOTAL POWER 1171.50 1035.0 Battery Capacity (A Hr) Allowed DoD Useable Capacity (A Hr) Time in Mode (minutes) Capacity Used Project Reserve SDO Preliminary Design Review (PDR) March 9-12, 2004 100 0.6 60

72 51.22 17.14% 100.0 0.6 60.0 72.0 47.2 27.05% Nov-03 Dec-03 Dec-03 Jan-04 Feb-04 Feb-04 Date Eclipse Mode Power Margin Percentage Margin Eclipse mode shows comfortable margin Eclipse Mode Power Performance Amp-Hours Subsystem Current Allocation Eclipse DKW 30.00% 25.00% 20.00% 15.00% Oct-03 Nov-03 Dec-03 Dec-03 Jan-04 Feb-04 Feb-04 Date Budget based on 100 A-hr battery with one failed cell. Eclipse is 72 minutes long and allowed DoD is 60%. Spacecraft Overview Page 37 SDO Pointing/Alignment Budgets Pointing allocations very similar to those at SCR, as AIA has similar requirements to the SPECTRE and Magritte instruments. Science reference absolute pointing (10, 3)) Heritage requirement from KCOR, kept as an internal requirement to serve as foundation for meeting all other requirements below High frequency: 5; seasonal variation 3; calibration errors 1: Total error: 9 (3)) Most significant contributor is spacecraft jitter; thermal shift expected to be <1, well within allocation

AIA absolute pointing (75, 3)) High frequency: 5; seasonal variation 4; biases and other one-time effects 48.75: Total error: 57.75 (3)) Most significant contributor are ground/launch shift effects Sun in Guide Telescope Linear FOV (95, 3)) Total error: 77.75 (3)) Worst-case assumption, based on AIA absolute pointing estimate summed with 20 SciTel/GT coalignment per AIA telescope Requirement set so that HMI does not have to adjust alignment legs on a daily basis (every two weeks is the current baseline) High frequency: 5; seasonal variation 5.92; leg adjustment errors 2.0: Total error: 12.92 (3)) Current analysis shows thermal shift allocated at 5 is less than 1 after an eclipse, leaving unofficial CBE near 8, well within requirements HMI in adjustment range of legs (200, 3)) Required to make sure all four AIA Guide Telescopes can be used simultaneously HMI in IMC field of view (14, 3)) Requirements baseline held off to better understand AIA requirements and performance, but rapidly converging on final allocations For the areas of most concern (AIA absolute pointing, HMI in IMC FOV), there is adequate margin for performance surprises High frequency: 5; seasonal variation 5.92; biases and other one-time effects 129.80: Total error: 140.72 (3)) Comfortable margin available to staying within range of legs EVE/ESP absolute pointing (450, 3)) High frequency: 5; seasonal variation 5.92; biases and other one-time effects 304.15: Total error: 315.07 (3)) Comfortable margin available to staying within EVEs performance range Jitter performance (monitored as project risk #58) described on next two charts Conservative HGA pointing budget shows 0.25 met with on-orbit calibration SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 38 Spacecraft Jitter Detailed discussions with HMI & AIA teams developed the two formulae that define the SDO Jitter requirement. HMI alone is concerned with misregistration, primarily concerned with disturbances between approximately 0.01 - 10 Hz (difficulty in combining filtergrams to create a Dopplergram) HMI and AIA are both concerned with blurring, all disturbances above 0.01 Hz are included Formulae below show the agreed method for calculating jitter at each focal plane PSD represents the disturbance PSD caused by spacecraft and instrument disturbances ATF is the attenuation transfer function of the HMI or AIA Image Stabilization System The final terms are frequency weighing functions, which result in the frequency ranges discussed above 2

RMS Misregistration = PSD( ) * ATF ( ) * 4 * sin c 2 ( ) * sin 2 ( ) d Figure1 Unattenuated PSD 10 10 10 10 10 10 10-2 10 -4 -8 10 10 -10 -1 -2 ATF G(s) -12 10 -2 10 -1 10 0 10 1 Frequency (Hz) 10 2 10 0 0 -2 -6 1/ 2 Cumulative Summation CCD Cumulative RMS (asec) 10 Attenuated PSD 0 2 CCD PSD(z) (asec /Hz) 10

2 RMS Blurring = PSD( ) * ATF ( ) * 1 sin c 2 ( ) d Figure 2 Attenuation Transfer Function Magnitude ) (asec2/Hz) Interface PSD( 10 1/ 2 10-4 10-6 10-8 10-10 10 10 10 -1 -2 -3 -4 10-12 3 10 -3 10 -5 -1 10 0 10 1 Frequency (Hz) SDO Preliminary Design Review (PDR) March 9-12, 2004 10 2 10 3 10-2 10-1 100 101 Frequency (Hz) 102 10 103 10-2 10 -1 10 0 10

1 10 2 10 Frequency (Hz) Spacecraft Overview Page 39 3 Spacecraft Jitter Jitter requirements met with margin if careful attention placed on RWA/structure interaction Analysis shows that of all disturbances considered, the major contributor is RWA imbalance resonating at a spacecraft flexible mode (HGA, GT NEA, torque quantization, slosh, instrument mechanisms also analyzed) Results assume four wheels at same speed, with commercially available mass balancing Results show that both ISS are adequate to effectively eliminate expected 0.01 - 10 Hz disturbances Table below shows the impact of limiting wheel speeds below 1800 rpm (30 Hz), most significant modes (besides HGA/solar array modes near 2 Hz) are above 30 Hz To establish additional jitter margin, ACS can operationally limit wheel speeds < 1800 rpm by ensuring that momentum buildup does not require a wide range of wheel speeds on-orbit (current estimate <1000 rpm) Category Requirement HMI Misregistration HMI Blurring AIA Blurring 0.03 0.14 0.17 Performance, wheels Performance, wheels not limited <1800 rpm 0.004 0.004 0.07 0.008 0.16 0.04 CCD Cumulative RMS (asec) -2 -4 10 -1 2 10 -6 10 -8 -10 -12 10 -2 10 -1 10 0

10 1 Frequency (Hz) 10 2 HGAS10-3resonance -10 10 -3 10 -12 10 10 -2 AIA, unlimited wheel SDO Preliminary Design Review (PDR) March 9-12, 2004 10 0 Frequency (Hz) 10 2 -2 10 -8 10 RWA resonance 10 -6 10 -2 -1 10 -4 10 10 10 10 -2 10 2 CCD PSD(z) (asec /Hz) /Hz) CCD PSD(z) (asec 10 CCD Cumulative RMS (asec) all values in arcseconds, 1 sigma -2 10 -1

10 0 10 1 10 2 10 Frequency (Hz) 3 10 -4 10 -2 10 -1 10 AIA, limited wheel 0 10 1 10 2 10 Frequency (Hz) Spacecraft Overview Page 40 3 10 High Gain Antenna Pointing POINTING ERROR BUDGET SDO High Gain Anntenna System Parameter (3 values) 4-Feb-04 Subsystem Requirement. A/T ACS ACS ACS A A A 0.10 Comm 1.00 0.13 0.13 HGAS/Mech HGAS/543 Comm/544 A/T T T T 0.14 0.14 Comm/544 Gimbal/544 T

T 7 Gnd (Deg) Req # 3.2.1 3.2.2 3.2.3 3.3.1 3.3.2 3.3.3 3.3.4 3.3.5 3.3.6 3.4.1 3.4.2 3.4.3 3.4.4 ACS/GN&C Knowledge/Command Errors ACS pointing knowledge Ephemeris accuracy Algorithm accuracy Hardware Ali gnment Errors Antenna boresight error Boom to S/C alignment error Gimbal to boom axis co-alignment error Gimbal to Gimbal-HGA I/f alignment error HGA to Gimbal-HGA I/f alignment error Gimbal interaxial orthogonality Launch/Deployment/Gravity Release Errors 1 Very Low Low Freq. High Freq. Freq. 6 Boom launch shift Boom gravity release 0.00 HGAS/543 A 0.10 HGAS/543 Boom deployment repeatability Gimbal actuator interface launch shifts 0.50 0.55 0.00 HGAS/543 Gimbal/544 Com A&T T T A 0.00 0.01 Com Mech. A A 0.01 Mech A 0.04 HGAS/ACS

0.02 0.02 0.08 ACS ACS Gimbal/544 A/T A A A 4 4 3.4.7 3.7.1 2 0.01 HGAS to S/C reference launch shift HGAS to S/C reference gravity release 3.6.1 3.6.2 3.6.3 3.6.4 (Deg) 8 0.01 Antenna launch shift Antenna gravity release 3.5.1 3.5.2 3.5.3 3.5.4 On-orbit n Error Static (Deg) Random (Deg.) 0.04 3.4.5 3.4.6 3.4.8 Calibratio Bias 4 Dynami c Pointi ng Errors Gimbal/boom dynamic interaction ACS induced boom dynamics Other S/C induced dynamics Gimbal tracking error Thermal Distortion Boom base to gimbal Gimbal, end-to-end Antenna S/C, S/C reference to HGADS I/F Dynami c Calibration Error 0.01 0.03 0.02 0.02 9 Column Totals (RSS), total error 0.05 1.04 0.75 HGAS/543 Gimbal

Com Mechanical Ground/ GN&C/RF no Cal 0.04 0.04 0.09 8 10 0.05 Column Totals (RSS), total error n-Cal Tot 0.04 0.04 0.09 (1) A/T - verification by analysis or test/inspection. (5) Test ETU in 1g; FLT by analysis/similarity only. (2) Measure and corrected after on-orbit deployment (6) If possible, test to verify dynamic interactions. (4) Pinned interface. (7) Bias errors assumed measurable, removable by compensation. (8) Calibration done right after Ephem. Upload. (9) Errors in calibration due to limits of calibration. (10) Multiple calibrations averaged to remove thermal distortion bias. SDO Preliminary Design Review (PDR) March 9-12, 2004 5 6 With on-orbit calibration, HGA pointing meets pointing requirement, even assuming conservative XTE/TRMM pointing assumptions A A A/T A A 0.92 0.23 Spacecraft Overview Page 41 Science Data Capture/Completeness/Bandwidth A simple way of expressing the HMI data capture needs is capture 99.99% of the data, 95% of the time Data capture: 95% requirement, generally across longer time scales includes outages over large time periods, due to eclipses, maneuvers, rain, etc., where multiple HMI observations are lost Data completeness: 99.99% requirement generally over short time scales includes dropouts (bit errors) over small time periods resulting in individual lost frames Multiple lost frames cause a single HMI observation to be lost Definition time scale is 10 minute observation period Up to 300 frames can be lost (of 3,000,000) in 10 minutes and meet data completeness When this number is exceeded, losses fall into data capture category Data capture also covers maneuvers, eclipses, etc where all data is lost SDO Data Capture is met with ample margin and detailed in SDO Data Capture Budget The data capture requirement for HMI has recently been changed to 95% over a 72-day period.

Twenty two (22) 72-day periods over five years must meet the 95% data capture requirement to meet full mission success. The 72-day periods will not be scheduled and nominally will run consecutively. If an data loss (or multiple losses) causes the 95% budget to be missed, a new 72-day period will be started A 72-day period centered around eclipse season meets the 95% data capture criteria with very little margin SDO Data Completeness is met by our RF link design and prevention of bit errors in instrument or spacecraft electronics Science data frames are compressed therefore one uncorrectable bit error results in one lost frame Frame error rate described by 99.99% can be converted to a bit error rate (BER) and allocations for BER distributed among subsystems in the science data path BER for hardware components is achieved by evaluating semiconductor parts for radiation tolerance and Single Event Effects (SEEs) rates; BER for RF downlink is part of the link calculation and is calculated after forward error correction such as Reed-Solomon and Convolutional encoding. Almost the full 130 Mbps science downlink capability is allocated to the instruments to avoid wasteful downlink of fill frames Individual instrument allocations regulated by Ka Comm card TDM table which samples data from six separation 1355 ports Replacement of SHARPP with AIA resulted in reallocation of data rate among instruments AIA originally requested 58 Mbps, now allocated 67 Mbps EVE originally requested 2 Mbps, now allocated 7 Mbps, eliminates compression altogether HMI already using lossless compression, constant at 55 Mbps SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 42 Data Capture/Completeness Breakouts Data Capture / Availability Budget for HMI, EVE, AIA Science Data Revision (A) Freq (per Duration yr) (hrs) HMI Hrs Lost /Year EVE Hrs Lost /Year AIA Hrs Lost /Year Notes: Operations 22.0 48.0 48.0 Includes Momentum Dumping,Stationkeeping, recovery periods, Antenna Handovers, Antenna Blockage 87.0 239.0 99.0 Includes Roll, Off-Point, EVE Flat Field, HMI Alignment, GT Cal, Internal Cals, CCD Outgassing, and recovery

104.7 104.7 194.7 Includes Eclipse, Lunar Eclipse, recovery, and radiation SEE's 112.5 112.5 112.5 Includes Maintenance, Rain, Solar RFI, Equipment Problems 0.8 31.6 7.9 Sub-total (Operations) Instrument/Science Calibrations/Activities Sub-total (Instrument Calibrations) Orbit Sub-total (Orbit) SDO Ground Station Sub-total (Ground Station) Completeness (incl. data loss due to Bit Errors from RF BER and SEE BER) random data loss due to BER (single bit flips) Totals HMI EVE AIA 327.03 535.76 462.09 95.00% 90.00% 90.00% 96.27% 93.89% 94.73% Planned/Allocated gaps - hrs/year Data Capture Reqt. Data Capture % Margin (hrs) 111.29 340.20 Margin (days) 4.64 14.18 Margin (%) 25.39% 38.81% Note: Margin will allow for data loss due to observatory and/or ground system anomalies On Board Electronics Instrument HMI 1.0E-10 AIA 5.0E-08 EVE 2.25E-8 414.53 17.27 47.29% RF Downlink Ka-Card RF Link Ground 1.0E-10 5.0E-09 0.0 Completeness reqt: HMI = 99.99%, EVE = 99.60%, AIA

= 99.90% Sum of all the planned/predicted outages/gaps MRD requirement for data availability/capture % of the year that data is available Difference between allowed outages and planned outages Difference between allowed outages and planned outages Ratio of unused/unallocated outages to allowed outages Completeness (Required) = 99.991% (99.990%) SDO Preliminary Design Review (PDR) March 9-12, 2004 = 99.921% (99.900%) = 99.638% (99.600%) Spacecraft Overview Page 43 Propellant Budget Overview Unlike the other technical resource budgets, the propellant budget does not lend itself to traditional margin calculations Of the propellant used for a mission, approximately 90% is a deterministic value dealing with the Delta V to get from GTO to GEO: only the last 10% has any potential for growth Propulsion has derived a margin requirement at 3% of tank capacity, as long as all worst case propellant drivers have been taken into account and are stacked together Among propellant growth contributors included 3200 kg separation mass Worst-case launch dispersions (11.2 m/s, 3 sigma) ISP variations in main engine or ACS thrusters (3% Isp penalty: 9 sec Isp reduction) Unexpectedly high pressure drop on either fuel or oxidizer lines, offsetting mixture ratio (37 kg penalty) Worst case ACS duty cycles and momentum unloading requirements Knowledge uncertainty, included in order to ensure there is adequate fuel for disposal when EOL is determined (currently budgeted at 5% of tank capacity, although estimates for value vary between 2-5%) Unusable residuals (1.15% penalty) For the ACS backup cases, immediate failure of main engine, replaced by ACS thrusters at lower ISP (19 sec Isp penalty) Main Engine, W-C Stack EELV dispersions GTO/GEO (ME +ACS) Stationkeeping Momentum Unloading Disposal Unusable Residual Knowledge Uncertainty Mass Ratio Losses TOTAL SDO Preliminary Design Review (PDR) March 9-12, 2004 Ox [kg] Fuel [kg] 7.5 773.1 0.8 8.7 5.5 10.1 46.3 17.5 869.5

4.5 468.4 0.5 5.2 3.3 6.0 28.0 18.9 534.8 %Ox of total %Fuel of total 0.9 0.8 88.9 87.6 0.1 0.1 1.0 1.0 0.6 0.6 1.2 1.1 5.3 5.2 2.0 3.5 100.0 100.0 Spacecraft Overview Page 44 Propellant Budget Status After incorporating all of the lessons learned from other missions and the Propulsion PDR and stacking worst-case propellant assumptions on top of each other, there is still positive propellant margin Assuming main engine use and worst-case stackup described before, there is 3.8% margin left in the tanks Nominal main engine case has 10.1% margin, when average dispersions and performance assumed Nominal ACS thruster backup has 5.0% margin, assuming immediate main engine failure but average performance of ACS thrusters and dispersions Realistically conservative ACS thruster backup case (which RSSs dispersions, 5% knowledge error, and fuel/oxidizer mass ratio losses) still has 1.8% margin Preliminary discussions with Nick Johnson (NASAs Orbit Debris Program Office) confirm that in this failure case, SDO will be able to use the knowledge uncertainty propellant to continue the mission, and thus may be able to gain back 5% margin Although SDO has an approach that meets our derived margin requirements, SDO will track a risk (# 73) against its propellant margin in order to protect the margin that exists 42 diameter tanks are used for propellant, next larger existing design is 49 in diameter, and would require redesign of the Observatory structure Although relatively minor (1-3%) gains are possible from continued optimization of the Propulsion design (e.g. higher Isp ACS thrusters), SDOs primary risk mitigation is to launch into a higher perigee GTO, and thus require less propellant A candidate perigee of 2000 km might coincide with EELV orbit debris requests, and would result in an additional 10% propellant margin Approach does not require redesign of the Observatory, but would require relaxation of Level 1 GTO orbit insertion requirements SEC customer has been briefed on this risk and is supportive of the mitigation if it is necessary SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 45 Launch Vehicle

Candidate Launch Vehicles: Delta IV Medium and Atlas V 401, with pre-launch processing at AstroTech As discussed in the SCR, SDO has followed KSC guidance and has designed with the possibility of either launch vehicle through PDR In order to mitigate design risk, SDO may rely on KSC (or GSFC) in-house analysis for each launch vehicle until candidate is selected KSC in-house capability for preliminary orbit analysis, particularly necessary in trading higher perigee options, preliminary assessment of launch constraints and describing operations/ground track timelines Several options (including GSFC in-house capability) being considered for getting Coupled Loads Analysis for each candidate vehicle L/V contract award based on Interface Requirements Document, scheduled for release by March 2004. KSC document with GSFC input Contract award targeted for July, 2004 Several detailed design questions await this procurement Per the Planners Guides, there is a static envelope interference with one of the vehicles Selection of PAF interface has a direct impact on PM design Tip-off rate performance for selected vehicle determines the likelihood of thruster firings after separation Final trajectory design, including launch constraints, may impact Launch power budget or operations concept Requirements on the mechanical design may be eased slightly by selection of one vendor (frequency requirements currently envelope both vendors, but they are not equal) Determination of vehicle opens discussion on other capabilities (communication access through vehicle, etc) SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 46 SDO Fairing Clearances 19.6 clearance to AIA (Aperture covers will be closed) 13.2 clearance to star trackers 12.6 clearance To star tracker 13.6 clearance to edge Of solar Array PAF = 1664-4 (65.5) Fairing = Composite, 4 m dia. 13.6 clearance To corner of S/A 18.5 clearance to AIA (Aperture covers will be closed) 12.0 from AIA Aperature Door Hinge

to the tapered section of fairing 17.3 inches clearance To HGA 17.3 Clearance to Antenna dish PAF Stay-out zone Vertical clearance = 4.1 To main engine Delta IV SDO Preliminary Design Review (PDR) March 9-12, 2004 Atlas 401 Spacecraft Overview Page 47 Delta IV Static Envelope Violation Static Envelope Delta IV 4 m Faring Static Envelope Sep. Plane Main Engine extends 4.58 in. (.12 m) below the LV static Envelope (Main Engine ~0.15 in. (0.38 cm) above PAF Diaphragm) PAF Diaphragm Delta IV PAF Note: 1) It is possible for SDO to extend the LV PAF interface ring to accommodate the Delta IV static envelope interference. Preliminary structural analysis shows that a 6 inch PAF extension does not have significant impact on fundamental modes and frequencies. The KSC ELV office has recommended for SDO to maintain its current design through the LV IRD development process. 2) The SDO Current design fits within Atlas V 4 m fairing static envelope (not shown). SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 48 Future Work Several tasks must continue to be worked as SDO moves from PDR to CDR: Continue to track development of breadboards, working towards early interface tests where appropriate (especially spacecraft/instrument interfaces) and starting life tests early where needed Complete procurement of spacecraft components so that detailed interfaces and expected performance is known at CDR, and any necessary detailed design adjustments can be made Once detailed component design is known, work Fault Protection detailed design, developing component-level detection and pairing it with mode and system-level tests Concept already exists for higher-level tests, but component-level cant be worked until components are known in order to develop correct data processing tests Continue to review jitter assumptions, to make sure all other contributions (HGA stepping, Guide Telescope NEA, instrument motion, among others) have defined requirements to make sure jitter can be met with assumed Reaction Wheel performance and operations Operational consideration for keeping wheels between areas of modal significance wont help if another disturbance is also a major contributor Subsystem teams have already built HGA command interface to allow for varying the period of HGA step commands, to make sure the HGA boom modes arent excited Continue to review propellant budget and determine if the risk mitigation must be triggered SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 49 Conclusions The SDO spacecraft meets all of its requirements with margin and is ready to move forward to detailed design

SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 50 Backup Charts SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 51 Formulation Trades Formulation Trades Trade Considered Orbit Selected Option Options Considered Inclined GEO L1; Sun-Synch Right Ascension of Ascending Orbit design Node = 200 EELV (either Delta IV or Atlas Launch vehicle Delta II (two or three stage) V) Horizontal mount (Triana Mechanical configuration SOHO-like vertical box design); Triangular optical bench Rationale continuous data downlink, HMI doppler range requirements Electrical data and power Distributed hierarchy system architecture Centralized services; Peer network allows for common building block design, develops subsystem "nodes" with simple interfaces to the rest of the spacecraft Subsystem data processor UTMC R000; 80196; Mongoose V In-house Ethernet design, inhouse Spacewire design tight regulation of PSE output for heaters Nickel Hydrogen short booms with daily handovers; long booms with full coverage Motorola Coldfire RH-CF5208 High-speed data interface 1355 using Atmel ASIC Power bus regulation for heaters follow normal 28 V heaters bus Battery chemistry Lithium Ion HGA boom configuration RF frequency range for science link KA transmitter output power/HGA size/boom length combined trade medium booms with tapered arrays minimizes eclipse seasons throughout mission provided necessary mass and volume to orbit provided clear thermal radiator field of view for all instruments, allowed growth for instruments, did not require an instrument stack, placed all IMC-based instruments on the same panel as their guide telescope no memory paging required, good throughput/power performance existence of a part that met data throughput needs without development, interface and part already used between Camera Electronics and Instrument Electronics cost, complexity and power loss of regulation design vs heater switching power/mass density; Lithium Ion appropriate for GEO orbit (few cycles) full mission coverage achievable via twice a year handovers or a 180 degree roll flip; preferrable from mass/flexibility standpoint to long booms, from an operations/data loss standpoint to short booms Ka band Ku, X only band with enough allowed bandwidth allocation to meet mission requirements

2.5 W transmitter, 0.75 m^2 dish, 1.7 m boom 5 W transmitter, 0.5, 1 m^2 dish, short or long boom best combination of acceptable RF link, mass, structural design HGA dish design dual reflector single reflector; slotted waveguide array less design complexity that array, less waveguide loss, smaller volume than single reflector Instrument module material composite aluminum to minimize thermal distortions through eclipse, during mission ACS science pointing use Guide Telescope only Star Trackers Flight Software development base EO-1/MAP heritage Triana/Swift; JWST UML development SDO Preliminary Design Review (PDR) March 9-12, 2004 direct measurement of target, better noise performance from GT than ST, one additional set of sensors to interface with familiarity of flight software development team, existing algorithms (may augment with Triana or JWST code) Spacecraft Overview Page 52 Detailed Mass Breakdown SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 53 Detailed Power Breakdown SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 54 Processor Resources Utilization Estimate Volatile Memory Non-Volatile Memory CPU Cycles Main Processor Requirements for Procurement 128 MB 4 MB 56 MIPS Main Processor 8 MB Apps 1.5 MB (X2 Copies) 20 MIPS Use Estimate 90 MB Recorder SDN Design Requirements 2 MB

2 MB 11 MIPS at 64 kb of Bootstrap 12 MHz PSE SDN Estimate based on MAP PSE RSN 110 Kbytes < 128 Kbytes 42% of 12 MIPS ACE SDN Estimate based on MAP ACE RSN 200 Kbytes < 512 Kbytes 20% of 12 MIPS S-Band SDN Estimate based on MAP Comm RSN 90 Kbytes < 128 Kbytes 11% of 12 MIPS GCE/HK SDN Estimate 150 Kbytes < 256 Kbytes 50% SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 55 1553 Bus Bandwidth Resource Utilization SDO GN&C Components AIA GT Data GCE HGA Position Cmds A/B GCE HGA Position Data A/B ACE Sensor Data A/B ACE SDN HK A/B ACE SDN Event Msg A/B ACE EVD-RWA Cmds A/B ACE EVD Tlm A/B ST Data A/B ST HK A/B HMI Image Stabilization Data GN&C Sub Total SDO Preliminary Design Review (PDR) March 9-12, 2004 Slots 5 10 10 10 2 2 10 10 10 2 1 72 SDO 1553 Slot Usage (4 ms Slots) Retry ? SDO 1553 Component Y S-Comm Code Blocks (2 kbps) A/B Y S-Comm RT Tlm (32 kbps*2) A/B Y S-Comm SDN HK A/B Y S-Comm SDN Event Msg A/B N PSE SDN HK A/B N PSE SDN Event Msg A/B

Y GCE/HK SDN HK A/B Y GCE/HK SDN Event Msg A/B Y Ka-Comm Card A/B N SBC RT Mode N Generic RT Commands HMI HK HMI Event Msg EVE HK EVE Event Msg AIA HK AIA Event Msg Instrument Diagnostic Tlm Time and Synch GN&C Sub Total Unused Total Used Total Available Slots 8 64 4 2 10 2 4 2 2 1 20 1 1 1 1 1 1 4 3 72 46 204 250 % 3% 26% 2% 1% 4% 1% 2% 1% 1% 0% 8% 0% 0% 0% 0% 0% 0% 2% 1% 29% 18% 82% Retry ? Y Y N N N N N N N N Y N N N N N N N N Spacecraft Overview Page 56

Detailed Pointing Breakdown (1 of 4) Requirement/ CBE (Y/Z) Allocation Calibrated Pointing of Science Reference Boresight 10.00 Description Responsibility Verification (from MRD 2.5.5.1, with 10" pointing necessary to meet various other budgets below) High frequency jitter requirement 5.00 High freq total 5.00 Seasonal Control Guide telescope error 1.00 ACS controller error 1.00 Drift of the SRB 1.41 CGT thermal shift 2.00 IM thermal shift (CGT to SRB) 1.00 Seasonal total 3.00 Separately budgeted allocations, found in SDO Jitter Budget. This number reflects specified jitter Difference between sun as imaged on GT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Thermal shift over the period of a month between AIA boresights and Science Reference Boresight, internal to telescopes. Only two telescopes define the SRB in each axis Each telescope may drift 2 arcsec Alignment shift between CGT and SRB over the period of a month, internal to CGT Alignment shift between the SRB and the mounting I/F for AIA/CGT that is being used, caused by IM deformation various, allocated separately AIA GT SC ACS AIA Telescope AIA GT SC MECH/THERM Bias Calibration error between SRB and CGT 1.00 Bias total 1.00 SRB Calibrated Pointing allocation 9.00 RSS Margin (Jitter) RSS Margin (Seasonal) RSS Margin (Bias)

3.32 2.65 1.73 Requirement/ CBE (Y/Z) Allocation How well science team can calibrate out alignment errors for SRB on-orbit, building in offset to GT's Description 95.00 For all four Guide Telescopes (calibrated pointing plus these effects must be less than 60", anticipated linear range of GT) jitter requirement 5.00 Separately budgeted allocations, found in SDO Jitter Budget. This number reflects specified jitter High freq total 5.00 Seasonal Control Guide telescope error 1.00 ACS controller error 1.00 CGT/mount I/F thermal shift 2.00 GT/mount I/F thermal shift 2.00 Sun in GT linear range AIA/SOC Responsibility High frequency IM thermal shift (GT to CGT) 1.00 Seasonal total 3.32 Difference between sun as imaged on GT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Alignment shift between CGT and Instrument Module over the period of a month, internal to CGT Alignment shift between GT and Instrument Module over the period of a month, internal to GT Alignment shift between CGT mounting I/F and mounting I/F for GT, caused by IM deformation various, allocated separately AIA GT SC ACS AIA GT Verification These budgets show that the reference boresight can be pointed to the sun with required accuracy, and that the coalignment of the Guide Telescopes is good enough to ensure overlap between the linear ranges of all of the Guide Telescopes

AIA GT SC MECH/THERM Bias GT to AIA Telescope alignment 20.00 AIA Absolute Pointing Bias total 75.00 77.62 AIA Reference Calibrated Pointing allocation 85.94 RSS Margin (Jitter) RSS Margin (Seasonal) RSS Margin (Bias) 14.14 12.90 40.65 SDO Preliminary Design Review (PDR) March 9-12, 2004 How well GT boresight is aligned with AIA telescope on the ground how stable the alignment is through environments See AIA Absolute Pointing Budget AIA/SC MECH AIA/SC MECH Spacecraft Overview Page 57 Detailed Pointing Breakdown (2 of 4) Requirement/ CBE (Y/Z) Allocation AIA Absolute Pointing (all other telescopes) 75.00 Description Responsibility (requirement is 75", based on acceptable offset of sun and heliosphere on CCD) High frequency jitter requirement 5.00 High freq total 5.00 Seasonal Control Guide telescope error 1.00 ACS controller error 1.00 Drift of the AIA 2.00 CGT/mount I/F thermal shift 2.00 GT/mount I/F thermal shift 2.00 IM thermal shift (CGT to SRB) 1.00 IM thermal shift (GT to SRB) 1.00 Seasonal total

4.00 Separately budgeted allocations, found in SDO Jitter Budget. This various, allocated number reflects specified jitter separately Difference between sun as imaged on GT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Thermal shift over the period of a month between SRB and Instrument Module, internal to the telescopes Alignment shift between CGT and Instrument Module over the period of a month, internal to CGT Alignment shift between GT and Instrument Module over the period of a month, internal to GT Alignment shift between SRB and mounting I/F for CGT, caused by IM deformation Alignment shift between SRB and mounting I/F for GT, caused by IM deformation AIA GT SC ACS AIA Telescope AIA GT AIA GT SC MECH/THERM SC MECH/THERM Bias Calibration error between SRB and CGT 1.00 Calibration error between SRB and GT 1.00 AIA telescope boresight knowledge 10.00 Integration Error of AIA Telescope to MRC 20.00 Mechanical shift in SRB due to AIA shifts 24.75 Mechanical shift in SRB due to internal AIA shifts 3.54 Mechanical shifts between AIA and mounting I/F 35.00 Mechanical shifts between AIA and mounting I/F due to internal AIA shifts 5.00 Bias total 48.75 AIA Reference Calibrated Pointing allocation 57.75 RSS Margin (Jitter) RSS Margin (Seasonal) RSS Margin (Bias) 21.68 20.87 44.49 SDO Preliminary Design Review (PDR) March 9-12, 2004 How well science team can calibrate out alignment errors for SRB on-orbit, building in offset to CGT How well science team can calibrate out alignment errors for SRB on-orbit, building in offset to GT's Measurement error in determining alignment of AIA telescope to its cube Using cube/cube alignment, how well AIA telescopes can be aligned with Master Reference Cube (MRC) on the groundhow parallel can the boresights be placed Change in alignment of the SRB due to launch, thermal settling, 1 g effects, etc Change in alignment of the SRB due to launch, thermal settling, etcmisalignments internal to instrument accounted for here Change in alignment of AIA to its mounting I/F due to launch,

thermal settling, 1 g effects, etc Change in alignment of AIA to its mounting I/F due to launch, thermal settling, etcmisalignments internal to instrument accounted for here AIA/SOC AIA/SOC AIA Telescope AIA/SC MECH SC MECH/THERM AIA Telescope SC MECH/THERM AIA Telescope Verification This budget shows that the AIA telescopes can be pointed to the sun with required accuracy. It assumes the reference boresight is pointed at the center of the sun, and allocates static shifts from the reference boresight and dynamic shifts from the Guide Telescope. Spacecraft Overview Page 58 Detailed Pointing Breakdown (3 of 4) Requirement/ CBE (Y/Z) Allocation HMI Calibrated Pointing High frequency 14.00 jitter requirement 5.00 High freq total 5.00 Seasonal Control Guide telescope error 1.00 ACS controller error 1.00 HMI/mounting I/F thermal shift 2.00 CGT/mount I/F thermal shift 2.00 IM thermal shift (CGT to HMI) 5.00 Seasonal total 5.92 Bias Adjustment error in HMI legs Bias total 2.00 2.00 AIA Reference Calibrated Pointing allocation 12.92 RSS Margin (Jitter) RSS Margin (Seasonal) RSS Margin (Bias)

3.47 3.74 2.35 HMI in range of legs High frequency Description Separately budgeted allocations, found in SDO Jitter Budget. This number reflects specified jitter 5.00 High freq total 5.00 Seasonal Control Guide telescope error 1.00 ACS controller error 1.00 HMI/mounting I/F thermal shift 2.00 CGT/mount I/F thermal shift 2.00 IM thermal shift (GT to HMI) 5.00 Seasonal total 5.92 SC MECH/THERM How well the HMI alignment can be corrected using adjustable legs HMI Separately budgeted allocations, found in SDO Jitter Budget. This number reflects specified jitter Difference between sun as imaged on CGT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Thermal shift over the period of a month between HMI boresight and Instrument Module, internal to telescope Alignment shift between CGT and Instrument Module over the period of a month, internal to CGT Alignment shift between HMI Optics Package mounting I/F and mounting I/F for CGT that is being used, caused by IM deformation AIA GT SC ACS HMI AIA GT Responsibility various, allocated separately AIA GT SC ACS HMI AIA GT SC MECH/THERM Bias HMI boresight knowledge 10.00 Coalignment of HMI CCDs 10.00 Integration of HMI to S/C,aligned to MRC 20.00 Mechanical shifts between HMI optical package and mounting

35.00 I/F Mechanical shift in SRB due to AIA shifts Mechanical shift in SRB due to internal AIA shifts 24.75 3.54 One-time mechanical shifts btw SRB and HMI mounting I/F120.00 Bias total AIA Reference Calibrated Pointing allocation RSS Margin (Jitter) Verification various, allocated separately Difference between sun as imaged on CGT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Thermal shirt over the period of a month between HMI boresight and Instrument Module, internal to HMI optics package Alignment shift between CGT and Instrument Module over the period of a month, internal to CGT Alignment shift between HMI Optics Package mounting I/F and mounting I/F for CGT that is being used, caused by IM deformation Requirement/ CBE (Y/Z) Description Allocation 200.00 (calibrated pointing plus these effects must be less than 200") jitter requirement Responsibility (requirement is 14", based on range of the HMI IMC system) Measurement error in determining alignment of HMI boresights to its optical package cube How coaligned are the two CCDs internal to the HMI package Using cube/cube alignment, how well HMI Optics Package can be aligned with the MRC on the groundhow parallel can the boresights be placed Change in alignment of HMI Optics Package/CCDs to its mounting I/F due to launch, thermal settling, 1 g effects, etcmisalignments internal to instruments accounted for here Change in alignment of the SRB due to launch, thermal settling, 1 g effects, etc Change in alignment of the SRB due to launch, thermal settling, etcmisalignments internal to instrument accounted for here Alignment shifts in Instrument Module slip due to mounting I/F is accounted for here HMI HMI AIA/SC MECH HMI Verification These budgets show HMI can be pointed to the sun with enough accuracy to stay within the adjustment range of its legs. Also, once HMI is adjusted, dynamic errors will not position the sun outside the field of view of HMIs Image Motion Compensation system. SC MECH/THERM AIA Telescope SC MECH 129.81 140.72 #VALUE! SDO Preliminary Design Review (PDR) March 9-12, 2004

Spacecraft Overview Page 59 Detailed Pointing Breakdown (4 of 4) Requirement/ CBE (Y/Z) Allocation EVE Absolute Pointing 450.00 Description Responsibility Verification (tightest requirement is ESP's 450", based on 3/21/03 Greg Ucker e-mail) High frequency jitter requirement 5.00 High freq total 5.00 Seasonal Guide telescope error 1.00 ACS controller error 1.00 EVE/mounting I/F thermal shift 2.00 GT/mount I/F thermal shift 2.00 IM thermal shift (GT to EVE) 5.00 Seasonal total 5.92 Separately budgeted allocations, found in SDO Jitter Budget. This various, allocated number reflects specified jitter separately Difference between sun as imaged on GT and output error signal The residual attitude error in the ACS controller, due to quantization, cyclic torque, etc Thermal shirt over the period of a month between EVE/ESP and Instrument Module, internal to EVE package Alignment shift between GT and Instrument Module over the period of a month, internal to GT Alignment shift between EVE package mounting I/F and mounting I/F GT that is being used, caused by IM deformation AIA GT SC ACS EVE AIA GT SC MECH/THERM Bias EVE (ESP) boresight knowledge WRT Reference Optical Sight 270.00 Integration of EVE to S/C, aligned to MRC 30.00 Mechanical shifts between EVEand mounting I/F 60.00 Mechanical shift in SRB due to AIA shifts 24.75 Mechanical shift in SRB due to internal AIA shifts 3.54

One-time mechanical shifts btw SRB and EVE mounting I/F 120.00 Bias total 304.01 AIA Reference Calibrated Pointing allocation 314.93 RSS Margin (Jitter) RSS Margin (Seasonal) RSS Margin (Bias) 139.98 140.86 316.81 Internal measurement error of ESP boresight, measured during EVE calibration at instrument level, with respect to optical sight on EVE Using cube/optocal flat alignment, how well EVE package can be aligned with MRC on the groundhow parallel can the boresights be placed Change in alignment of EVE to its mounting I/F due to launch, thermal settling, etcmisalignments internal to instrument accounted for here Change in alignment of the SRB due to launch, thermal settling, 1 g effects, etc Change in alignment of the SRB due to launch, thermal settling, etcmisalignments internal to instrument accounted for here Alignment shifts in Instrument Module SRB and EVE packageslip due to mounting I/F is accounted for here EVE EVE/MECH EVE SC MECH/THERM AIA Telescope SC MECH This budget shows the tightest EVE pointing requirement is met by SDO. SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 60 Jitter Sensitivity to RWA Speed The most significant jitter contribution is wheel imbalance at a structural resonance 0 10 Acceptable RWA range 10 (jitter considerations only) Against a rigid body, wheel imbalance causes <0.01 of jitter, but at worst case significant modes, imbalance can result in 0.014 of jitter Significant structural modes in the 2-4 Hz range and above 30 Hz PSD (asec2/Hz) Lower frequency disturbances such as fuel slosh and HGAS pointing attenuated by HMI and AIA IMC systems -2 -4 10 -6 10 -8 10 10

10 10 10 10 -10 10 -2 Unattenuated PSD -4 0 10 1 10 2 10 -6 -8 -12 10 Wheel speed (in Hz) vs disturbance Attenuated PSD -10 10-2 -2 10 -1 10 0 10 1 10 2 10 3 Frequency (Hz) 10-4 10 0 10 -1 10 -2 10 -3 10-6 10-8 10-10 10-12 10-2 10-1 100

101 Frequency (Hz) CCD Cumulative RMS (asec) Interface PSD() (asec2/Hz) 10 0 2 /Hz) CCD PSD(z) (asec 10 102 10 The plot above shows the peak magnitude to be expected on a PSD when a wheel runs a a specific spee. Only one of these points will be on a PSD at any given time. If the wheel can be kept in an acceptable range, then the disturbance can be mitigated. As one can see in the accumulated jitter plot lower right, almost 0.1RMS jitter is created from the spike in the PSD, lower left/center. If wed stayed in the lower frequency zone, wed seen 1/10 of the disturbance. (HMI plots shown.) 103 -4 -5 10 -2 10 10 -1 10 0 10 1 Frequency (Hz) SDO Preliminary Design Review (PDR) March 9-12, 2004 10 2 10 3 Cummulative Summation Spacecraft Overview Page 61 Details of Data Capture Budget SDO Preliminary Design Review (PDR) March 9-12, 2004 Spacecraft Overview Page 62 Propellant Budget Details MAIN ENGINE NOMINAL CASE OXIDIZER FUEL TOTAL [kg] [kg] [kg] EELV dispersions 0 0 0.0 A1 253.7 153.8 407.4 A2 246.8 149.5

396.3 A3 225.5 136.7 362.2 A4 31 18.8 49.8 A5 0.7 0.4 1.1 A6 0.1 0.0 0.1 A7 0.1 0.0 0.1 A8 N/A N/A N/A A9 N/A N/A N/A A10 N/A N/A N/A A11 N/A N/A N/A A12 N/A N/A N/A A13 N/A N/A N/A A14 N/A N/A N/A A15 N/A N/A N/A A16 N/A N/A N/A A17 N/A N/A N/A A18 N/A N/A N/A A19 N/A N/A N/A A20 N/A N/A N/A EWSK 0.8 0.5 1.3 Momentum Unload 8.8 5.3 14.2 Disposal 5.6 3.4 8.9 MR outage propellant 0 0 0 Knowledge Error 33.9 19.8 53.7 Tank Residuals 9.5

5.7 15.2 Manifold Residuals 0.6 0.3 0.9 CATEGORY TOTAL [kg] TANK CAPACITY Margin to capacity [kg] Margin to capacity [%] 817 909 91.9 10.1 494 MAIN ENGINE WORST CASE ACS THRUSTER NOMINAL CASE ACS THRUSTER WORST CASE OXIDIZER FUEL TOTAL OXIDIZER FUEL TOTAL OXIDIZER FUEL TOTAL [kg] [kg] [kg] [kg] [kg] [kg] [kg] [kg] [kg] 7.5 4.5 12.0 0.0 0.0 0.0 8.0 4.9 12.9 260 157.6 417.6 54.8 33.2 88 56.2 34.1 90.3 251.8 152.6 404.4 49.3 29.9 79.2 50.6 30.7 81.3 229 138.8 367.8 50.1 30.4 80.5 51.3 31.1 82.4 31.4 19 50.4 54.7 33.2 87.9 56.0 34.0 90.0 0.7 0.4 1.1 53 32.1 85.1 54.2 32.9 87.1 0.1

0 0.1 54.7 33.1 87.8 55.9 33.9 89.7 0.1 0 0.1 54.8 33.2 88 55.9 33.9 89.8 N/A N/A N/A 53.7 32.6 86.3 54.8 33.2 88.0 N/A N/A N/A 49.4 30 79.4 50.4 30.5 80.9 N/A N/A N/A 49.8 30.2 80 50.7 30.7 81.4 N/A N/A N/A 54.2 32.9 87.1 55.1 33.4 88.5 N/A N/A N/A 49.2 29.8 79 50.0 30.3 80.3 N/A N/A N/A 49.8 30.3 80.1 50.5 30.6 81.1 N/A N/A N/A 53.5 32.4 85.9 54.2 32.9 87.1 N/A N/A N/A 47.2 28.6 75.8 47.8 29.0 76.8 N/A N/A N/A 12.1 7.3 19.4

12.2 7.4 19.6 N/A N/A N/A 7.8 4.7 12.5 7.9 4.8 12.7 N/A N/A N/A 0.7 0.4 1.1 0.7 0.4 1.1 N/A N/A N/A 0.3 0.2 0.5 0.3 0.2 0.4 N/A N/A N/A 0.1 0.1 0.2 0.1 0.1 0.2 0.8 0.5 1.3 1.1 0.6 1.6 1.0 0.6 1.6 8.7 5.2 13.9 12 7.3 19.2 11.9 7.2 19.1 5.5 3.3 8.8 7.5 4.6 12.1 7.5 4.5 12.0 17.5 18.9 36.4 0 0 0 18.1 19.6 37.7 46.3 28.0 74.3 33.9 19.8 53.7 46.3 28.0 74.3 9.5 5.7 15.2 9.5 5.7 15.2 9.5 5.7 15.2 0.6 0.3

0.9 0.6 0.3 0.9 0.6 0.3 0.9 NOTE: Errors RSS'ed 1311 870 535 1404 864 523 1387 896 547 1442 556 kg 61.8 kg 11.1 % SDO Preliminary Design Review (PDR) March 9-12, 2004 909 39.5 4.3 556 kg 21.2 kg 3.8 % 909 45.2 5.0 556 kg 33.1 kg 6.0 % 909 13.3 1.5 556 kg 9.3 kg 1.7 % Spacecraft Overview Page 63

Recently Viewed Presentations

  • Budowa komputera osobistego (PC)

    Budowa komputera osobistego (PC)

    Obecnie wielkość tej pamięci to średnio 8 MB (jeszcze do niedawna przeciętna pamięć wynosiła 512 Kb), a coraz częściej 16 lub 32 Mb. W pamięci tej przechowywane są dane o każdym punkcie obrazu, a także tekstury (w postaci map bitowych)...
  • Modern American poetry - Mrs. Croswell&#x27;s Classroom

    Modern American poetry - Mrs. Croswell's Classroom

    Imagists and Imagist Poetry. Imagists - group of writers who revolted against forms of the past & desired to write poetry for aesthetic purposes. Aesthetic - pleasing to the eye. ... Modern American poetry Last modified by: JJ Croswell ...
  • STUDY ABROAD ORIENTATION For Summer &amp; Fall 2010

    STUDY ABROAD ORIENTATION For Summer & Fall 2010

    SAFETY ABROAD. Safety, both at home and abroad, are hot topics these days. Some catastrophic events—the recent wildfires in Chile, flooding in the Southern US, terrorist attacks in London get lots of attention--and are largely out of anyone's control. However,...
  • Cal Poly State University, SLO HVAC Best Practi

    Cal Poly State University, SLO HVAC Best Practi

    CSU Sustainability Policy . Major revision to policy since last CSU Executive Order 987 (2006) Academics: "The CSU will seek to further integrate sustainability into the academic curriculum working within the normal campus consultative process."
  • 3 GA Efroymson Viability PR - Conservation Gateway

    3 GA Efroymson Viability PR - Conservation Gateway

    The "Total Score" column shows numerical values (0-10) that help describe the magnitude of an overall threat rank at a finer level than is given by the Very High, High, Medium, and Low ranks. These scores are used for creating...
  • Chapter 3

    Chapter 3

    The 3 Pillars of OOP&D. Object-Oriented Programming. Encapsulation. Hide details in a class, provide methods. Polymorphism. Same name, different behavior, based on type. Inheritance. Capture common attributes and behaviors in a base class and extend it for different types
  • Diapositive 1 - saint-gobain-gyproc.com

    Diapositive 1 - saint-gobain-gyproc.com

    The New Academic Educational Building (NAEB) for the Royal College of Surgeons in Ireland builds on the College's heritage of excellence and innovation.
  • The Wars of Religion

    The Wars of Religion

    Catholic League led by Ferdinand (now HRE Ferdinand II), Maximilian of Bavaria & aided by Spanish . Battle of White Mountain (1620) - ends Bohemian phase. Frederick lost his lands in Palatinate to Spain. Bohemia given back to Ferdinand confiscate...