SDO Systems RetreatSolar Dynamics Observatory (SDO)
Mission PDR
David Ward
Agenda
Spacecraft trades and major changes since SCR
Overview of architecture and spacecraft subsystems
Observatory operations concept
Issues
Conclusion
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 Spring’s procurement activities
At this point, spacecraft and subsystem requirements are in good shape to proceed into detailed design
Spacecraft Overview Page *
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
Spacecraft Overview Page *
Design Trades Performed Since SCR
Sheet1
Resource
Project holds 20% reserve, rest allocated to instruments and subsystems
Power
25%
31.5%
sunlit mode, all instruments on, assumes lower voltage due to one failed cell
Eclipse Mode
Launch Mode
Orbit Injection Mode
Stationkeeping Mode
Survival Mode
36.2%
margin against 80% DoD, assumes fault at exit of eclsipe, followed by 30 minute recovery (also one failed batttery cell)
Survival Mode
Propellant
3% of total propellant
both cases include -3 sigma Isp, maximum ACS control and momentum buildup, worst case mixture ratio, disposal propellant
Main engine
CPU Throughput
based on specified SBC MIPS performance, thought achievable in industry
SDN's
58%
assumes 12 MIPS processing, which may require oscillator change from prototype
1553 Bus Bandwidth
S Downlink on Subcarrier
assumes 32 kbps downlink to SDO or commercial Ground Station
Formulation Trades
Trade Considered
Selected Option
Options Considered
Orbit design
minimizes eclipse seasons throughout mission
Launch vehicle
Delta II (two or three stage)
provided necessary mass and volume to orbit
Mechanical configuration
Horizontal mount (Triana design); Triangular optical bench
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
Electrical data and power system architecture
Distributed hierarchy
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
Motorola Coldfire RH-CF5208
no memory paging required, good throughput/power performance
High-speed data interface
In-house Ethernet design, in-house Spacewire design
existence of a part that met data throughput needs without development, interface and part already used between Camera Electronics and Instrument Electronics
Power bus regulation for heaters
heaters follow normal 28 V bus
tight regulation of PSE output for heaters
cost, complexity and power loss of regulation design vs heater switching
Battery chemistry
Lithium Ion
Nickel Hydrogen
power/mass density; Lithium Ion appropriate for GEO orbit (few cycles)
HGA boom configuration
short booms with daily handovers; long booms with full coverage
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
RF frequency range for science link
Ka band
Ku, X
only band with enough allowed bandwidth allocation to meet mission requirements
KA transmitter output power/HGA size/boom length combined trade
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
single reflector; slotted waveguide array
less design complexity that array, less waveguide loss, smaller volume than single reflector
Instrument module material
ACS science pointing
use Guide Telescope
only Star Trackers
direct measurement of target, better noise performance from GT than ST, one additional set of sensors to interface with
Flight Software development base
Design Trades
Trade Considered
Selected Option
Options Considered
separation perigee at 300 km
nominal insertion at 185 km
minimizes momentum buildup during GTO phase, allowing for use of available Reaction Wheels without requiring thruster control during perigee passes
Propulsion module design
thrusters fire in three different directions, four tank module
stationkeeping can be performed with thrusters only in one direction, allowed for Propulsion Module to completely be a separate unit from spacecraft module, for parallel I&T; two tank stack eased development of PM, moved slosh mass onto centerline
Propulsion design
Traditional MMH/NTO biprop
Dual-mode N2H4/NTO biprop/monoprop
once all thrusters anti-sunward, contamination concerns from MMH significantly reduced; less complex design; higher Isp of ACS htrusters allows for a backup to orbit in event main engine fails
Propulsion Module thermal design
Hybrid of "toasty cavity" around tanks and internal and individually wrapped thruster lines
traditional design with individual wrapping; only a toast cavity
simplifies heater design around tanks, reduces risk of a problem with an unaccessible heater/thermostat, uses an acceptable amount of heater power compared to toasty cavity alone
Solar array sizing
7.7 m^2
5.8 m^2 and PSE modifications
small savings in solar array size did not outweigh cost/complexity/risk of abandoning MAP heritage DET design
Omni antenna location
Omni boresights angled between X and Z in XZ plane
either +/- X or +/- Z boresights
X axis boresights have antenna pattern nulls located such that stationkeeping burns would be affected, Z axis boresights would place nulls during eclipse
Ka transmitter design: integrated vs componentized
Integrated design
Componentized (components connected via coax)
baseline design is integrated transmitter, for best noise performance and ability to trade-off performance among various components as more is learned during design
Reaction wheel make/buy
Out-of-house design
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
Number and location of AIA guide telescopes
Four AIA Guide Telescopes. Each assigned to one Science Telescope
Two AIA Guide Telescopes, shared among the four AIA Science Telescopes
eliminates concern about local flexibility between GuideTels and SciTels, makes jitter attenuation design very similar to TRACE
EVE science data bus allocation
7 Mbps
2 Mbps
AIA science data bus allocation
67 Mbps
58 Mbps
HMI science data bus allocation
55 Mbps
Sheet2
Sheet3
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
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
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/A’s autonomously deployed at separation, HGA’s deployed individually by ground command after Observatory is power-positive
Increased Solar Array size without affecting HGA coverage
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 EVE’s 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
Spacecraft Overview Page *
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 FOV’s (more difficult if either had KCOR’s 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
POST SCR Config. 8/03
Design Changes: New Thruster Locations
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 *
Observatory Mechanical Configuration
Two short tapered arrays, cell side out when stowed
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)
AIA (which uses GT signals for IMC) all on one face, HMI, EVE (which do not) on the other
EVE
AIA
EVE
HMI
AIA
Spacecraft bus module provides Faraday cage and radiation shielding for s/c and instrument components
HMI
SDO Propulsion Module
SDO Spacecraft Subsystem Overview
Hardware command decoding for computer-free recovery
Provides continuous 130 Mbps high-speed interface between Instruments and Ka-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
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
Attitude Control
Jitter performance at focal plane to <0.5” (3σ), calibrated pointing accuracy of 10” (3σ) via zero-momentum, three-axis control with Reaction Wheels
Star Tracker, Inertial Reference Unit, and Guide Telescope used for target/attitude determination
Momentum unloading monthly with thrusters
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
Thermostatic control of survival heaters
Hybrid approach or “toasty cavity” and individual line heaters to minimize risk in propulsion thermal design
Spacecraft Overview Page *
SDO Electrical Architecture
Thermistors,
actuators & heaters
DC-DC Converter
Ka
DC-DC Converter
DC/DC Converter
High Speed
Common Design Elements
Subsystem Data Node
Serves as the embedded processor for many of the spacecraft avionics boxes
Uses a Motorola RH-CF5208 ColdFire processor for processing
Provides MIL-STD-1553 for communications with the spacecraft processor and cPCI for backplane communications
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
Subsystem Power Node
Since SCR, split into two boards to provide all of the common power requirements
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 LPSC’s are on switched 28V services)
Spacecraft Overview Page *
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
Spacecraft Overview Page *
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
Instrument Commissioning Phase
Science Mission Phase
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
Spacecraft Overview Page *
Launch and Acquisition
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 15º 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
Spacecraft Overview Page *
Launch and First Orbit Timeline
Spacecraft Overview Page *
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
Inertial Hold / Slew Capability checkout prior to fist planned maneuver
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 3rd 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
Spacecraft Overview Page *
Orbit Circularization
Phase used during first weeks to circularize the orbit from the GTO and place SDO in its geosynchronous slot at 102ºW.
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
Spacecraft Overview Page *
In-Orbit Checkout & Orbit Circularization
In-Orbit Checkout & Orbit Circularization
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
Spacecraft Overview Page *
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
Spacecraft Overview Page *
Nominal Mission Phase (Science Phase)
Spacecraft Overview Page *
Periodic Calibrations/Housekeeping
There are several periodic interruptions to the nominal science mission mode:
Stationkeeping and Momentum Unload maneuvers, Instrument Calibration (Roll and Off-point) 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
Spacecraft Overview Page *
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
Spacecraft Overview Page *
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
Spacecraft Overview Page *
Eclipse Timeline
Stationkeeping/Momentum Management
Required operations to keep SDO within its orbit “slot” at 102º 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 +/- 5º
During Delta-V spacecraft may be off pointed up to 45º from the sun-line for up to 30 minutes
Maximum offset of 15º from the XY plane to minimize sun on instrument CCD radiators
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.
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
Spacecraft Overview Page *
Typical Stationkeeping Timeline
Spacecraft Overview Page *
Safehold/Emergency Modes
Several capabilities will exist on the Observatory for “safing” in the event of an anomaly
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 “I’m OK” communications between ACE and s/c processor will cause ACE to enter Safehold
Safehold works without ground intervention until momentum limits reached
Momentum capabilities sized for one week of control before ground commanded unload performed
Safehold will not autonomously fire thrusters to unload, avoiding possibility of tumbling spacecraft
Safehold and Sun Acq are sun pointing. Same orientation as nominal science pointing
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 LPSC’s (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)
Critical event notification commands have been identified to inform instruments of current or pending conditions.
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
Spacecraft Overview Page *
Disposal
At end of mission, NASA policy requires disposal of SDO into an orbit that won’t 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 cross-sectional 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
Spacecraft Overview Page *
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, SDO’s 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”
Spacecraft Overview Page *
Technical Resources Status/Margins
Detailed breakdowns of mass, power, 1553 budgets available in backup charts.
Sheet1
Resource
SCR
PDR
CDR
Flight
Project holds 20% reserve, rest allocated to instruments and subsystems
Power
30%
25%
15%
30.0%
sunlit mode, all instruments on, assumes lower voltage due to one failed cell
Eclipse Mode
27.0%
margin against 60% DoD, assumes one failed batttery cell, 72 minutes, normal power configuration
Launch Mode
28.3%
margin against 60% DoD, assumes one failed batttery cell, assumes 120 for entire phase, 45 minutes launch to separation
Orbit Injection Mode
25.3%
margin against 60% DoD, assumes one failed batttery cell, 90 minutes entire phase, 50 minute burn
Stationkeeping Mode
139.9%
margin against 60% DoD, assumes one failed batttery cell, 45 degrees off sunline
Survival Mode
34.3%
margin against 80% DoD, assumes fault at exit of eclsipe, followed by 30 minute recovery (also one failed batttery cell)
Survival Mode
Propellant
positive margin with 3 sigma usage stackup
worst-case stackup includes -3% Isp, maximum ACS control and momentum buildup, worst case mixture ratio, disposal propellant
Main engine
ACS thruster backup
5.0%
propellant margin at 3200 kg, nominal performance values, except Isp for ACS thrusters during Orbit Injection
Main engine, w-c stack
ACS b/u, RSS stack
Memory
50%
50%
40%
25%
CPU Throughput
based on specified SBC MIPS performance, thought achievable in industry
SDN's
58%
assumes 12 MIPS processing, which may require oscillator change from prototype
1553 Bus Bandwidth
S Downlink on Subcarrier
assumes 32 kbps downlink to SDO or commercial Ground Station
Formulation Trades
Trade Considered
Selected Option
Options Considered
Orbit design
minimizes eclipse seasons throughout mission
Launch vehicle
Delta II (two or three stage)
provided necessary mass and volume to orbit
Mechanical configuration
Horizontal mount (Triana design); Triangular optical bench
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
Electrical data and power system architecture
Distributed hierarchy
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
Motorola Coldfire RH-CF5208
no memory paging required, good throughput/power performance
High-speed data interface
In-house Ethernet design, in-house Spacewire design
existence of a part that met data throughput needs without development, interface and part already used between Camera Electronics and Instrument Electronics
Power bus regulation for heaters
heaters follow normal 28 V bus
tight regulation of PSE output for heaters
cost, complexity and power loss of regulation design vs heater switching
Battery chemistry
Lithium Ion
Nickel Hydrogen
power/mass density; Lithium Ion appropriate for GEO orbit (few cycles)
HGA boom configuration
short booms with daily handovers; long booms with full coverage
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
RF frequency range for science link
Ka band
Ku, X
only band with enough allowed bandwidth allocation to meet mission requirements
KA transmitter output power/HGA size/boom length combined trade
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
single reflector; slotted waveguide array
less design complexity that array, less waveguide loss, smaller volume than single reflector
Instrument module material
ACS science pointing
use Guide Telescope
only Star Trackers
direct measurement of target, better noise performance from GT than ST, one additional set of sensors to interface with
Flight Software development base
Design Trades
Trade Considered
Selected Option
Options Considered
separation perigee at 300 km
nominal insertion at 185 km
minimizes momentum buildup during GTO phase, allowing for use of available Reaction Wheels without requiring thruster control during perigee passes
Propulsion module layout
thrusters fire in three different directions, four tank module
stationkeeping can be performed with thrusters only in one direction, allowed for Propulsion Module to completely be a separate unit from spacecraft module, for parallel I&T; two tank stack eased development of PM, moved slosh mass onto centerline
Propulsion design
Traditional MMH/NTO biprop
Dual-mode N2H4/NTO biprop/monoprop
once all thrusters anti-sunward, contamination concerns from MMH significantly reduced; less complex design; higher Isp of ACS htrusters allows for a backup to orbit in event main engine fails
Propulsion Module thermal design
Hybrid of "toasty cavity" around tanks and internal and individually wrapped thruster lines
traditional design with individual wrapping; only a toast cavity
simplifies heater design around tanks, reduces risk of a problem with an unaccessible heater/thermostat, uses an acceptable amount of heater power compared to toasty cavity alone
Solar array sizing
7.7 m^2
5.8 m^2 and PSE modifications
small savings in solar array size did not outweigh cost/complexity/risk of abandoning MAP heritage DET design
Omni antenna location
Omni boresights angled between X and Z in XZ plane
either +/- X or +/- Z boresights
X axis boresights have antenna pattern nulls located such that stationkeeping burns would be affected, Z axis boresights would place nulls during eclipse
Ka transmitter design: integrated vs componentized
Integrated design
Componentized (separate components connected via coax)
baseline design is integrated transmitter, for best noise performance and ability to trade-off performance among various components as more is learned during design
Reaction wheel make/buy
Out-of-house design
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
Number and location of AIA guide telescopes
Four AIA Guide Telescopes. Each assigned to one Science Telescope
Two AIA Guide Telescopes, shared among the four AIA Science Telescopes
eliminates concern about local flexibility between GuideTels and SciTels, makes jitter attenuation design very similar to TRACE
Reallocation of science downlink rate after SHARPP removal: EVE allocation
7 Mbps
2 Mbps
Reallocation of science downlink rate after SHARPP removal: AIA allocation
69 Mbps
58 Mbps
Reallocation of science downlink rate after SHARPP removal: HMI allocation
55 Mbps
Sheet2
Sheet3
Mass Budget Details
Mass budget shows comfortable margin
Budget based on 3200 kg separation mass. Margin measured against dry mass, given a known propellant load
cover
Mass Worksheet
Prepared by: D. Ward/594
CHECK THE SDO MIS AT https://sdoweb.gsfc.nasa.gov
TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.
&R464-SYS-SPEC-0008 Revision (-)
Change Record
DOCUMENT DATE: November 17, 2003
REVISION
DATE
1/30/04
Revision Date
Mass/Unit
Status
Notes
Mechanical
542.34
values based on updated Mech estimate, 10/24, which didn't separate fasteners
Propulsion Module
HGA Booms
Zeolite Filter (accounted in PM)
0
0
5
Est
Power
96.50
PSE
42
1
42.00
Est
Battery
42.5
1
42.50
ETU
Solar Array Cells/Cover Glass
ACS
97.30
Attitude Control Electronics
number based on MIMU envelope, with 3 TARAs less mass
Reaction Wheels
Oxidizer Tanks
0.32
2
0.16
ETU
0.4
2
0.20
Est
Fasteners
4.2
210
0.02
Est
C&DH
Thermal
45.30
MLI
31.90
1
31.90
Est
OSR's
0.00
0
1.00
Est
AEB
22.51
1
22.51
Est
Telescopes
72.04
4
18.01
Est
ESP
3.27
1
3.27
EEB
14.96
1
14.96
6.6
1
6.6
Est/Ana
Updated at 10/03 project status review, based on 95% fill
Hydrazine
532.00
821
Ana
Oxidizer
878.00
589
Ana
Pressurant
5.00
5
Ana
Isp
g
Mo
Mf
Lnof
DV
210
9.805
1100
1000
0.0953101798
196.2484257261
405.95
29.44%
1790
147
210
9.805
0.0713921469
1.0740023099
132.4641347875
1922.4641347875
19-Jul-02
30.4
Upated C&DH estimate. Increased battery, S/A and PSE to reflect power growth. Doubled S-band for redundancy.
Included more information WRT propulsion, taken from 6/26 trade presentation. Added 12 CSS (guess).
26-Jul-02
25.39
Increased RWA to 14 kg/wheel; increased HGA to 30 kg; increased ST to 7.26 (RDOS), decreased IRU (red MIMU)
02-Aug-02
18
Increased HGA estimate to 80 kg (from 30) per C Monroe e-mail.
09-Aug-02
19.1
17-Aug-02
24.01
Instruments selected, HGA estimate dropped to 31/antenna, added 2 kg guess for HGA control
20-Aug-02
24.22
22-Oct-02
22.9
Updated estimate of propulsion mass, taking 4 tanks into account
01-Nov-02
3.6
Giulio's updates to structure (mass doubled), more detailed breakout, removed IM, HGA mass increased
04-Nov-02
8.15
Broke out instrument margin, added column for Isp assumption, corrected number of gimbals
07-Nov-02
9.04
Updates from Mike Powers on RF mass, removal of PiVoT
13-Nov-02
21.32
Fuel mass scrubbed, real numbers for HGA and structure; SHEB mass cut by 40%
27-Nov-02
20.31
10-Jan-03
17.9
10-Jan-03
17.8
04-Feb-03
33.6
Change to Delta IV 4040, limited mass to 2500 kg, added 2% margin to fuel, scaled up fuel and mechanical, added back in SHARPP elex mass
08-Feb-03
29.6
Added back Instrument margin from proposals, changed mechanical add-up to reach 12.5% of total mass, added some mass for harness, thermal control
18-Mar-03
29.49
Switched from 80 Ahr to 100 Ahr battery, added requested margin for EVE, HMI, Ka card included in CDH box estimate, updated prop and HGA estimates
(Growth in prop due to larger fuel mass, switch to diaphragm N2H4 tank), updated allocation to 300 kg
24-Mar-03
30.28
Received update from mechanical, increased allocation to 3200 kg, added ziolite filter
27-Mar-03
30.03
02-Apr-03
Corrected misallocation of CDH, adding 0.1 kg, propellant budget scrubbed, allocation remains the same
01-Jun-03
31.02
01-Jul-03
30.82
Increase HMI allocation by 2 kg, estimated split evenly between optics and electronics, per CCR 16
22-Jul-03
33.4
Update to EVE, SHARPP estimates based on CSRs, update to propulsion estimate based on GNCC review, update to mech estimate based on mech analysis
01-Aug-03
28.33
Update to harness estimate (110 to 137) based on discussions with Paul Kim, update to solar array estimate based on discussions with Jason Hair
08-Aug-03
37.27
Update to new MMH propellant, oxidizer, dry mass estimates, assuming enough mass for backup operations
08-Sep-03
37.27
04-Oct-03
39.21
13-Nov-03
31.8
Reassigned mass on the basis of CCR 30, with updates as described in early November 2003
05-Jan-04
31.03
29-Jan-04
30.32
10-Feb-04
29.77
24-Feb-04
29.44
Updated PM estimate, moved adsorber into PM, updated AIA estimate.
Mass Trend
DKW
0.97
0.97
0.97
1.01
1.01
1.02
1.02
1.02
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
Power
103.00
96.50
118.50
80.50
118.50
118.50
118.50
118.50
133.30
133.30
134.50
96.50
96.50
96.50
96.50
300.40
405.95
415.80
431.74
415.80
424.68
424.68
449.10
496.20
535.83
432.20
423.29
421.31
415.31
402.36
Contamination Rremoved
David Ward: In November, CCR30 reallocated mass from power, contamination, and mechanisms to mechanical. New baseline values shown below
Observatory Plots
Observatory Plots
Mass Worksheet
Prepared by: D. Ward/594
CHECK THE SDO MIS AT https://sdoweb.gsfc.nasa.gov
TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.
&R464-SYS-SPEC-0008 Revision (-)
Change Record
DOCUMENT DATE: November 17, 2003
REVISION
DATE
1/30/04
Revision Date
Mass/Unit
Status
Notes
Mechanical
542.34
values based on updated Mech estimate, 10/24, which didn't separate fasteners
Propulsion Module
HGA Booms
Zeolite Filter (included in PM)
0
1
0
Est
Power
96.50
PSE
42
1
42.00
Est
Battery
42.5
1
42.50
ETU
Solar Array Cells/Cover Glass
ACS
97.30
Attitude Control Electronics
number based on MIMU envelope, with 3 TARAs less mass
Reaction Wheels
Oxidizer Tanks
0.32
2
0.16
ETU
0.4
2
0.20
Est
Fasteners
4.2
210
0.02
Est
C&DH
Thermal
45.30
MLI
31.90
1
31.90
Est
OSR's
0.00
0
1.00
Est
AEB
22.54
1
22.54
Est/ETU
Telescopes
72.04
4
18.01
Est
ESP
3.27
1
3.27
EEB
14.96
1
14.96
6.6
1
6.6
Est/Ana
210
9.805
1100
1000
0.0953101798
196.2484257261
405.92
29.43%
1790
147
210
9.805
0.0713921469
1.0740023099
132.4641347875
1922.4641347875
19-Jul-02
30.4
Upated C&DH estimate. Increased battery, S/A and PSE to reflect power growth. Doubled S-band for redundancy.
Included more information WRT propulsion, taken from 6/26 trade presentation. Added 12 CSS (guess).
26-Jul-02
25.39
Increased RWA to 14 kg/wheel; increased HGA to 30 kg; increased ST to 7.26 (RDOS), decreased IRU (red MIMU)
02-Aug-02
18
Increased HGA estimate to 80 kg (from 30) per C Monroe e-mail.
09-Aug-02
19.1
17-Aug-02
24.01
Instruments selected, HGA estimate dropped to 31/antenna, added 2 kg guess for HGA control
20-Aug-02
24.22
22-Oct-02
22.9
Updated estimate of propulsion mass, taking 4 tanks into account
01-Nov-02
3.6
Giulio's updates to structure (mass doubled), more detailed breakout, removed IM, HGA mass increased
04-Nov-02
8.15
Broke out instrument margin, added column for Isp assumption, corrected number of gimbals
07-Nov-02
9.04
Updates from Mike Powers on RF mass, removal of PiVoT
13-Nov-02
21.32
Fuel mass scrubbed, real numbers for HGA and structure; SHEB mass cut by 40%
27-Nov-02
20.31
10-Jan-03
17.9
10-Jan-03
17.8
04-Feb-03
33.6
Change to Delta IV 4040, limited mass to 2500 kg, added 2% margin to fuel, scaled up fuel and mechanical, added back in SHARPP elex mass
08-Feb-03
29.6
Added back Instrument margin from proposals, changed mechanical add-up to reach 12.5% of total mass, added some mass for harness, thermal control
18-Mar-03
29.49
Switched from 80 Ahr to 100 Ahr battery, added requested margin for EVE, HMI, Ka card included in CDH box estimate, updated prop and HGA estimates
(Growth in prop due to larger fuel mass, switch to diaphragm N2H4 tank), updated allocation to 300 kg
24-Mar-03
30.28
Received update from mechanical, increased allocation to 3200 kg, added ziolite filter
27-Mar-03
30.03
02-Apr-03
Corrected misallocation of CDH, adding 0.1 kg, propellant budget scrubbed, allocation remains the same
01-Jun-03
31.02
01-Jul-03
30.82
Increase HMI allocation by 2 kg, estimated split evenly between optics and electronics, per CCR 16
22-Jul-03
33.4
Update to EVE, SHARPP estimates based on CSRs, update to propulsion estimate based on GNCC review, update to mech estimate based on mech analysis
01-Aug-03
28.33
Update to harness estimate (110 to 137) based on discussions with Paul Kim, update to solar array estimate based on discussions with Jason Hair
08-Aug-03
37.27
Update to new MMH propellant, oxidizer, dry mass estimates, assuming enough mass for backup operations
08-Sep-03
37.27
04-Oct-03
39.21
13-Nov-03
31.8
Reassigned mass on the basis of CCR 30, with updates as described in early November 2003
05-Jan-04
31.03
29-Jan-04
30.32
10-Feb-04
29.77
24-Feb-04
29.43
DKW
0.97
0.97
0.97
1.01
1.01
1.02
1.02
1.02
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
Power
103.00
96.50
118.50
80.50
118.50
118.50
118.50
118.50
133.30
133.30
134.50
96.50
96.50
96.50
96.50
300.40
405.92
420.74
451.60
420.74
429.64
429.64
454.10
446.70
480.90
437.20
428.29
421.31
415.31
405.92
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
Contamination Rremoved
David Ward: In November, CCR30 reallocated mass from power, contamination, and mechanisms to mechanical. New baseline values shown below
Observatory Plots
Observatory Plots
Mass Worksheet
Prepared by: D. Ward/594
CHECK THE SDO MIS AT https://sdoweb.gsfc.nasa.gov
TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.
&R464-SYS-SPEC-0008 Revision (-)
Change Record
DOCUMENT DATE: November 17, 2003
REVISION
DATE
1/30/04
Revision Date
Mass/Unit
Status
Notes
Mechanical
542.34
values based on updated Mech estimate, 10/24, which didn't separate fasteners
Propulsion Module
HGA Booms
Zeolite Filter (included in PM)
0
1
0
Est
Power
96.50
PSE
42
1
42.00
Est
Battery
42.5
1
42.50
ETU
Solar Array Cells/Cover Glass
ACS
97.30
Attitude Control Electronics
number based on MIMU envelope, with 3 TARAs less mass
Reaction Wheels
Oxidizer Tanks
0.32
2
0.16
ETU
0.4
2
0.20
Est
Fasteners
4.2
210
0.02
Est
C&DH
Thermal
45.30
MLI
31.90
1
31.90
Est
OSR's
0.00
0
1.00
Est
AEB
22.54
1
22.54
Est/ETU
Telescopes
72.04
4
18.01
Est
ESP
3.27
1
3.27
EEB
14.96
1
14.96
6.6
1
6.6
Est/Ana
210
9.805
1100
1000
0.0953101798
196.2484257261
405.92
29.43%
1790
147
210
9.805
0.0713921469
1.0740023099
132.4641347875
1922.4641347875
19-Jul-02
30.4
Upated C&DH estimate. Increased battery, S/A and PSE to reflect power growth. Doubled S-band for redundancy.
Included more information WRT propulsion, taken from 6/26 trade presentation. Added 12 CSS (guess).
26-Jul-02
25.39
Increased RWA to 14 kg/wheel; increased HGA to 30 kg; increased ST to 7.26 (RDOS), decreased IRU (red MIMU)
02-Aug-02
18
Increased HGA estimate to 80 kg (from 30) per C Monroe e-mail.
09-Aug-02
19.1
17-Aug-02
24.01
Instruments selected, HGA estimate dropped to 31/antenna, added 2 kg guess for HGA control
20-Aug-02
24.22
22-Oct-02
22.9
Updated estimate of propulsion mass, taking 4 tanks into account
01-Nov-02
3.6
Giulio's updates to structure (mass doubled), more detailed breakout, removed IM, HGA mass increased
04-Nov-02
8.15
Broke out instrument margin, added column for Isp assumption, corrected number of gimbals
07-Nov-02
9.04
Updates from Mike Powers on RF mass, removal of PiVoT
13-Nov-02
21.32
Fuel mass scrubbed, real numbers for HGA and structure; SHEB mass cut by 40%
27-Nov-02
20.31
10-Jan-03
17.9
10-Jan-03
17.8
04-Feb-03
33.6
Change to Delta IV 4040, limited mass to 2500 kg, added 2% margin to fuel, scaled up fuel and mechanical, added back in SHARPP elex mass
08-Feb-03
29.6
Added back Instrument margin from proposals, changed mechanical add-up to reach 12.5% of total mass, added some mass for harness, thermal control
18-Mar-03
29.49
Switched from 80 Ahr to 100 Ahr battery, added requested margin for EVE, HMI, Ka card included in CDH box estimate, updated prop and HGA estimates
(Growth in prop due to larger fuel mass, switch to diaphragm N2H4 tank), updated allocation to 300 kg
24-Mar-03
30.28
Received update from mechanical, increased allocation to 3200 kg, added ziolite filter
27-Mar-03
30.03
02-Apr-03
Corrected misallocation of CDH, adding 0.1 kg, propellant budget scrubbed, allocation remains the same
01-Jun-03
31.02
01-Jul-03
30.82
Increase HMI allocation by 2 kg, estimated split evenly between optics and electronics, per CCR 16
22-Jul-03
33.4
Update to EVE, SHARPP estimates based on CSRs, update to propulsion estimate based on GNCC review, update to mech estimate based on mech analysis
01-Aug-03
28.33
Update to harness estimate (110 to 137) based on discussions with Paul Kim, update to solar array estimate based on discussions with Jason Hair
08-Aug-03
37.27
Update to new MMH propellant, oxidizer, dry mass estimates, assuming enough mass for backup operations
08-Sep-03
37.27
04-Oct-03
39.21
13-Nov-03
31.8
Reassigned mass on the basis of CCR 30, with updates as described in early November 2003
05-Jan-04
31.03
29-Jan-04
30.32
10-Feb-04
29.77
24-Feb-04
29.43
DKW
0.97
0.97
0.97
1.01
1.01
1.02
1.02
1.02
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
Power
103.00
96.50
118.50
80.50
118.50
118.50
118.50
118.50
133.30
133.30
134.50
96.50
96.50
96.50
96.50
300.40
405.92
420.74
451.60
420.74
429.64
429.64
454.10
446.70
480.90
437.20
428.29
421.31
415.31
405.92
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
0.00
Contamination Rremoved
David Ward: In November, CCR30 reallocated mass from power, contamination, and mechanisms to mechanical. New baseline values shown below
Observatory Plots
Observatory Plots
Sunlit Power Budget Details
Normal mode shows comfortable margin
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.)
cover
Power Worksheet
Prepared by: D. Ward/594
CHECK THE SDO MIS AT https://sdoweb.gsfc.nasa.gov
TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.
&R464-SYS-SPEC-0008 Revision (-)
Change Record
DOCUMENT DATE: November 17, 2003
REVISION
DATE
23-Feb-04
&R464-SYS-SPEC-0008 Revision A
Mode Descriptions
MODE DESCRIPTIONS
The following worksheets allocate power for several Observatory Modes, as a way of describing the power allocations for various configurations. Here are brief summaries of the different modes:
SUNLIT/NORMAL MODE: This is meant to encompass the operational science-data-taking mode for each subsystem. In order to conservatively size the solar arrays, the configuration is taken durin eclipse season, with the battery recharging, but the battery v
ECLIPSE MODE: This configuration captures the Observatory as it passes through a normal eclipse. As of 11/13/03, the plan is to leave the instruments powered on through every eclipse, but they may be asked to conserve power by limiting the science outpu
LAUNCH MODE (not yet configured): This mode captures the variable power state of different components by time averaging the amount of time a component is on as compared to the total time of the mode (120 minutes). This is accomplished by multiplying the
ORBIT RAISING MODE (not yet configured): This mode captures the configuration of the Observatory during the long GTO burns to raise perigee. As with other modes, time averaging is performed in the number of components field to model the amount of time s
STATIONKEEPING MODE (not yet configured): This mode captures the short burns whle in GEO to maintain the orbit slot. Momentum unloads are similar, but do not require a possible slew 45 degrees off the sunline, which is modeled here. This sheets combine
SURVIVAL MODE (not yet configured): This configuration verifies the ability to survive a power or attitude anomaly at eclipse exit, and not get back power-positive for 30 minutes (102 minutes without power). In order to analyze a true worst-case, nomina
SURVIVAL MODE--STEADY STATE (not yet configured): This configuration verifies the ability to stay power-positive in the Survival Mode configuration
Detailed Data
0.5
42.2
0.5
38.6
0.5
26.3
0.5
37.4
0.5
52.6
0.5
38.6
0.5
31.7
Electrical
21.1
19.3
13.2
18.7
26.3
19.3
15.8
0.5
22.8
0.1
3.9
0.1
2.7
0.1
3.8
0.5
28.1
0.1
3.9
0.5
17.4
1121.3
1035.0
705.9
1001.7
1378.8
958.4
857.4
1121.3
1035.0
705.9
1001.7
1378.8
958.4
857.4
Derived capability
Solar array capacity, EOL (W) (per D Keys end of eclipse, w/o failures)
1653.4813533616
Solar array capacity, EOL (W) (per D Keys solar array model, eclipse, cell failure)
1457.9683554979
91.1
162.0
765.9
107.2
1408.2
47.46%
30.02%
27.05%
28.34%
25.28%
139.86%
34.34%
64.25%
19-Jul-02
24-Jul-02
17-Aug-02
Updated APS ST, SSPA, PSE and battery trickle charge numbers
27-Nov-02
9-Feb-03
26-Mar-03
Added allocations, considering revising thermal so that all numbers in one spot
2-Apr-03
20-May-03
Totaled Load Bus Power and then derived Array requirement and margin
27-May-03
29-May-03
Updated spreadsheet to reflect no 1.5 V drop during eclipse and other battery modes, per Mike/Denney comments
1-Jul-03
Added 1 W to HMI in normal and eclipse mode (and increased normal mode allocation) per CCR 16
11-Jul-03
Added simple solar array model and failure cases (cell failure, PWM failure, 21 V with cell failure)
8-Aug-03
Changed solar array sizing tool to reflect better computation of array power, different peak power point, better thermal coefficient
15-Sep-03
Rearranged harness to be calculated from load current, incorporated Denney's newest system sizing tool
24-Sep-03
Replaced SHARPP with LMSAL AIA estimates, used "Camera Read Out" (w/o margin) for normal mode
24-Oct-03
Updates from project status reviews, removed non-existant Cat Bed Htrs, Ka feed, battery voltages are average over event (eclipse/launch/burn), with one failed cell
26-Nov-03
Some formatting additions, addition of several trend worksheets, added survival and stationkeeping modes
28-Nov-03
Added power from solar arrays during slews in Orbit Raising and Survival
5-Jan-04
27-Jan-04
Updated per M Powers email, back to 53W, based on better estimate of efficiency and DC/DC power required
30-Jan-04
Updated to reflect AIA w/4 Guide Telescopes and EVE with survival and decon htrs on in Orbit Inject
10-Feb-04
Updated power to reflect HGAS growth (nominal case, assumes 65% efficiency)
11-Feb-04
Updates per February monthlies, powered off IM bench heaters during orbit raising to save power, added "steady-state" survival case
21-Feb-04
Update per CCR 55, ACE at 40.1, RWA at 18/, CDH @ 95, mech at 40.5
David Ward: this number based on 10 W/wheel 50% of time, 30 W/wheel 50% of time
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
David Ward: per Dave Steinfeld, 10/24/03 analysis
David Ward: per D Keys 2/04 PSR
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: should unheated option be considered to save power here?
David Ward: should unheated option be considered to save power here?
David Ward: estimte for average power, based on low slew rates, from Rich 2/10/04
David Ward: numbers updated perD Nguyen, 2/04
David Ward: derived from 38W parasitics, EPS PDR, 12/19/03
David Ward: per M Bay and D Keys, summer 2003
David Ward: per M Bay and D Keys, summer 2003
David Ward: per D Keys, 10/27/03, 2.5A is constant, no matter what voltage
David Ward: requested 10/24/03 by P Gonzales, allows for T-15 minutes umbilical separation, 60 minutes from launch to separation, 45 minutes to acquire
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
dward: 4 thrusters on for maximum of 50 minutes, compared to total maneuver
dward: main engine on for maximum of 50 minutes, compared to total maneuver
dward: HGA not planned for deployment at separation
dward: transmitter turned on ~15 minutes before separation, approx 60 min before acquisition is complete
dward: per B DeFazio longest eclipse at SCR
dward: set at voltage to create 24.5 at array/battery level
dward: four wheels on for 45 minutes, compared to total period which includes launch and pre launch
dward: estimated at steady state power for top wheel speed
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: numbers updated per J McCabe 10/03 PSR
David Ward: estimte for average power, based on low slew rates, from Rich 10/24/03
David Ward: numbers updated per E Grob 10/03 PSR
dward: set at voltage to create 24.5 at array/battery level
David Ward: per D Keys, 10/27/03
David Ward: per M Bay and D Keys, summer 2003
David Ward: per M Bay and D Keys, summer 2003
David Ward: per D Keys, 10/27/03, 2.5A is constant, no matter what voltage
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp…heaters powered off at survival
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per B DeFazio longest eclipse at SCR, plus 30 minutes
dward: voltage drop per M Bay, agreed to by P Kim, G Won in Elec Sys Spec
David Ward: 1 recovery slews, assume 15 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
David Ward: 2 slews, assume 10 of the 20 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
David Ward: 30 minutes for sun acquisition, assume 15 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
dward: assumed survival mode power for launch
dward: assumed early ops mode power for orbit inject
dward: survival heater plus decontamination heaters
David Ward: numbers updated per J McCabe 10/03 PSR
David Ward: numbers updated per J McCabe 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
dward: per CCR 55
dward: thermal power values are the average of cold case and survival case, per 2/04 D Nguyen
David Ward: per D Keys 2/04 PSR
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp…heaters powered off at survival
dward: thermal power values are the average of cold case and survival case, per 2/04 D Nguyen
dward: steady state power margin in survival, includes assumption of 15 dgrees offpointing
dward: assume 15 degrees off-pointing
dward: powered off to save power
dward: set at voltage to create 24.5 at array/battery level
David Ward: derived from 38W parasitics, EPS PDR, 12/19/03
David Ward: per M Bay and D Keys, summer 2003
David Ward: per M Bay and D Keys, summer 2003
David Ward: per D Keys, 10/27/03, 2.5A is constant, no matter what voltage
dward: assume 45 degrees off sunline
dward: per operations plan
dward: per CCR 55
dward: per CCR 55
dward: per CCR 55
dward: per CCR 55
dward: per CCR 55
dward: per CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
dward: from Rich,CCR 55
Normal Mode Trend Data
SDO Power by Subsystem
Started capturing faults
(eclipse exit, failed PWM)
Subsystem
Components
0.5
17.1
0.5
19.5
0.5
17.1
0.5
22.8
0.1
3.7
0.1
4.2
0.1
3.7
0.1
4.9
1085.0
1084.3
1085.0
1086.3
1085.0
1084.3
1085.0
1086.3
Volts
Amps
Watts
Volts
Amps
Watts
Volts
Amps
Watts
Volts
Amps
Watts
1653.5
1458.0
1488.1
1318.3
52.39%
34.47%
37.15%
21.36%
563/D. Keys
Initial Cell Efficiency 27.5% (This efficiency is at 135.3 mw/cm2)
0.275
0.99
(Watts)
Harness
0.99985
0.903
23.610%
S/C Loads
Array Panel Area Available (SDO Baseline area is 7.5 m^2)
7.5
Electronics load (Watts)
Load Input Voltage
Number of strings (baseline is 130), 18 series cells per string
130
717.3
29.6
String Ouput Voltage
String diode drop
Harness voltage drop
1653.5
28
28
Volts
1458.0
24.5
1488.1
28
Notes:
Input Parameters for S/C Load Value and Desired Operating Voltage only (Highlighted boxes)
100% Battery DOD Case (nominal)
1430.0
24
1318.3
22
Note:
Assumptions:
Must provide sufficient voltage headroom to charge LiIon battery to 4.2V/Cell (33.6V Max)
By design, use of boost regulator in PSE allows operation over knee of curve (constant voltage side of curve)
Worst case average (hot) temperature estimated by thermal group as 63 C
Worst case average (cold) temperature estimated by thermal group at -130 C
Thermal Correction Factors for S/A
Temperature
73
63
50
-130
-150
984.1
1325
1068.4
1325
1069.4
1325
1107.3
1438.1
1107.3
1438.1
1020.8
1458
1035.2
1458
1031.5
1458
1048.2
1458
1070.1
1458
1121.4
1458
cover
Power Worksheet
Prepared by: D. Ward/594
CHECK THE SDO MIS AT https://sdoweb.gsfc.nasa.gov
TO VERIFY THAT THIS IS THE CORRECT VERSION PRIOR TO USE.
&R464-SYS-SPEC-0008 Revision (-)
Change Record
DOCUMENT DATE: November 17, 2003
REVISION
DATE
23-Feb-04
&R464-SYS-SPEC-0008 Revision A
Mode Descriptions
MODE DESCRIPTIONS
The following worksheets allocate power for several Observatory Modes, as a way of describing the power allocations for various configurations. Here are brief summaries of the different modes:
SUNLIT/NORMAL MODE: This is meant to encompass the operational science-data-taking mode for each subsystem. In order to conservatively size the solar arrays, the configuration is taken durin eclipse season, with the battery recharging, but the battery v
ECLIPSE MODE: This configuration captures the Observatory as it passes through a normal eclipse. As of 11/13/03, the plan is to leave the instruments powered on through every eclipse, but they may be asked to conserve power by limiting the science outpu
LAUNCH MODE (not yet configured): This mode captures the variable power state of different components by time averaging the amount of time a component is on as compared to the total time of the mode (120 minutes). This is accomplished by multiplying the
ORBIT RAISING MODE (not yet configured): This mode captures the configuration of the Observatory during the long GTO burns to raise perigee. As with other modes, time averaging is performed in the number of components field to model the amount of time s
STATIONKEEPING MODE (not yet configured): This mode captures the short burns whle in GEO to maintain the orbit slot. Momentum unloads are similar, but do not require a possible slew 45 degrees off the sunline, which is modeled here. This sheets combine
SURVIVAL MODE (not yet configured): This configuration verifies the ability to survive a power or attitude anomaly at eclipse exit, and not get back power-positive for 30 minutes (102 minutes without power). In order to analyze a true worst-case, nomina
SURVIVAL MODE--STEADY STATE (not yet configured): This configuration verifies the ability to stay power-positive in the Survival Mode configuration
Detailed Data
0.5
42.2
0.5
38.6
0.5
26.3
0.5
37.4
0.5
52.6
0.5
38.6
0.5
31.7
Electrical
21.1
19.3
13.2
18.7
26.3
19.3
15.8
0.5
22.8
0.1
3.9
0.1
2.7
0.1
3.8
0.5
28.1
0.1
3.9
0.5
17.4
1121.3
1035.0
705.9
1001.7
1378.8
958.4
857.4
1121.3
1035.0
705.9
1001.7
1378.8
958.4
857.4
Derived capability
Solar array capacity, EOL (W) (per D Keys end of eclipse, w/o failures)
1653.4813533616
Solar array capacity, EOL (W) (per D Keys solar array model, eclipse, cell failure)
1457.9683554979
91.1
162.0
765.9
107.2
1408.2
47.46%
30.02%
27.05%
28.34%
25.28%
139.86%
34.34%
64.25%
19-Jul-02
24-Jul-02
17-Aug-02
Updated APS ST, SSPA, PSE and battery trickle charge numbers
27-Nov-02
9-Feb-03
26-Mar-03
Added allocations, considering revising thermal so that all numbers in one spot
2-Apr-03
20-May-03
Totaled Load Bus Power and then derived Array requirement and margin
27-May-03
29-May-03
Updated spreadsheet to reflect no 1.5 V drop during eclipse and other battery modes, per Mike/Denney comments
1-Jul-03
Added 1 W to HMI in normal and eclipse mode (and increased normal mode allocation) per CCR 16
11-Jul-03
Added simple solar array model and failure cases (cell failure, PWM failure, 21 V with cell failure)
8-Aug-03
Changed solar array sizing tool to reflect better computation of array power, different peak power point, better thermal coefficient
15-Sep-03
Rearranged harness to be calculated from load current, incorporated Denney's newest system sizing tool
24-Sep-03
Replaced SHARPP with LMSAL AIA estimates, used "Camera Read Out" (w/o margin) for normal mode
24-Oct-03
Updates from project status reviews, removed non-existant Cat Bed Htrs, Ka feed, battery voltages are average over event (eclipse/launch/burn), with one failed cell
26-Nov-03
Some formatting additions, addition of several trend worksheets, added survival and stationkeeping modes
28-Nov-03
Added power from solar arrays during slews in Orbit Raising and Survival
5-Jan-04
27-Jan-04
Updated per M Powers email, back to 53W, based on better estimate of efficiency and DC/DC power required
30-Jan-04
Updated to reflect AIA w/4 Guide Telescopes and EVE with survival and decon htrs on in Orbit Inject
10-Feb-04
Updated power to reflect HGAS growth (nominal case, assumes 65% efficiency)
11-Feb-04
Updates per February monthlies, powered off IM bench heaters during orbit raising to save power, added "steady-state" survival case
21-Feb-04
Update per CCR 55, ACE at 40.1, RWA at 18/, CDH @ 95, mech at 40.5
David Ward: this number based on 10 W/wheel 50% of time, 30 W/wheel 50% of time
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
David Ward: per Dave Steinfeld, 10/24/03 analysis
David Ward: per D Keys 2/04 PSR
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: should unheated option be considered to save power here?
David Ward: should unheated option be considered to save power here?
David Ward: estimte for average power, based on low slew rates, from Rich 2/10/04
David Ward: numbers updated perD Nguyen, 2/04
David Ward: derived from 38W parasitics, EPS PDR, 12/19/03
David Ward: per M Bay and D Keys, summer 2003
David Ward: per M Bay and D Keys, summer 2003
David Ward: per D Keys, 10/27/03, 2.5A is constant, no matter what voltage
David Ward: requested 10/24/03 by P Gonzales, allows for T-15 minutes umbilical separation, 60 minutes from launch to separation, 45 minutes to acquire
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
dward: 4 thrusters on for maximum of 50 minutes, compared to total maneuver
dward: main engine on for maximum of 50 minutes, compared to total maneuver
dward: HGA not planned for deployment at separation
dward: transmitter turned on ~15 minutes before separation, approx 60 min before acquisition is complete
dward: per B DeFazio longest eclipse at SCR
dward: set at voltage to create 24.5 at array/battery level
dward: four wheels on for 45 minutes, compared to total period which includes launch and pre launch
dward: estimated at steady state power for top wheel speed
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp
David Ward: numbers updated per J McCabe 10/03 PSR
David Ward: estimte for average power, based on low slew rates, from Rich 10/24/03
David Ward: numbers updated per E Grob 10/03 PSR
dward: set at voltage to create 24.5 at array/battery level
David Ward: per D Keys, 10/27/03
David Ward: per M Bay and D Keys, summer 2003
David Ward: per M Bay and D Keys, summer 2003
David Ward: per D Keys, 10/27/03, 2.5A is constant, no matter what voltage
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp…heaters powered off at survival
David Ward: one cell failed, set so that battery is average of 24.5 and 28V
dward: per D Keys, 10/27/03
dward: per D Keys, 10/27/03
dward: per B DeFazio longest eclipse at SCR, plus 30 minutes
dward: voltage drop per M Bay, agreed to by P Kim, G Won in Elec Sys Spec
David Ward: 1 recovery slews, assume 15 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
David Ward: 2 slews, assume 10 of the 20 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
David Ward: 30 minutes for sun acquisition, assume 15 minutes within 90 degrees of sunline, multiplied by integrated area under cosine curve
dward: assumed survival mode power for launch
dward: assumed early ops mode power for orbit inject
dward: survival heater plus decontamination heaters
David Ward: numbers updated per J McCabe 10/03 PSR
David Ward: numbers updated per J McCabe 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
David Ward: per D Keys 2/04 PSR
dward: per CCR 55
dward: thermal power values are the average of cold case and survival case, per 2/04 D Nguyen
David Ward: per D Keys 2/04 PSR
David Ward: TARA power with heater on, assuming duty cycle for coldest op temp…heaters powered off at survival
dward: thermal power values are the average of cold case and survival case, per 2/04 D Nguyen
dward: steady state power margin in survival, includes assumption of 15 dgrees offpointing
dward: assume 15 degree