1. SDO Electrical Power Subsystem (EPS) Mission PDR
Denney Keys
EPS Subsystem Lead
Mike Burns – Northrop-Grumman, PSE Lead
Ed Gaddy – Code 563, Solar Array Lead
Leo Lee – Code 563, Battery Lead
Dave Sullivan – Code 563, Battery GSE Lead
Gopal Rao – Code 563, Battery Support
Tom Rozanski – Code 563, Battery & Test Support
John Washington – Code 563, Cable & GSE Support

2. Driving Requirements Solar Array Battery
>1450 watts between 26.1 volts and 35.2 volts
< 11.9 kg of electrical equipment on the two wings
< 4A per circuit
> 5 year life meeting the specs above in a geosynchronous orbit
temperature between -170C and 79.9C for operating cells and 90.8C for shunted cells
reliability tbd (computed at .995)
outgassing requirements
no sparkling
no more than 50 volts between adjacent cells
ITO coating on covers (grounded conductive covers)
Battery
Maximum height of the battery must be < 11 inches
Maximum mass allocation for the battery is 42.5 Kg
Must operate at 29 +/- 6 V with >100 Ah capacity
Must operate for 5 years under GEO with a maximum DoD of 50% or 60% with one cell failure
The battery must be capable of cell by-pass and cell balancing

3. Driving Requirements Power Subsystem Electronics (PSE)
Main Spacecraft Power Bus Regulation
Process or shunt Solar Array power
1380W Orbit Ave Power (1500W peak)
Regulate Battery Dominated bus between 23V- 35V
Battery Charge Regulation
Regulate Commanded battery voltage to 100mV
Provide up to 10A max battery charge current
Power Distribution
Provide switched or un-switched power feeds to all spacecraft loads
Provide over-current protection on all feeds
Deployment Functions
Sense launch vehicle separation signal (2 out of 3)
Provide power command to actuate Solar Array deployment
Provide power command to actuate HGA deployment
Reliability
Single fault tolerant
Operational Life
5yrs
Radiation
100kRad Total Dose, 100MeV SEL and SEGR, 37Mev SEU

4. SDO EPS Descriptive Overview
SDO Power System consists of three primary components
Power System Electronics (PSE)
Solar Array (2 Panels)
Li-Ion Battery
Sunlight converted to electrical energy by Solar Array (Direct Energy Transfer)
Array power processed by PSE using digital and PWM shunts and distributed to spacecraft loads (through switched and unswitched outputs) and battery
Li-Ion battery used to store energy during solstice periods and provide full spacecraft power during eclipse periods
Power system architecture designed for single fault tolerance
8 for 7 battery cell redundancy
18 for 20 solar array circuit redundancy
Redundant backplane and control architecture utilized in PSE
PSE architecture heritage uses Microwave Anisotropy Probe (MAP) as basis for design
Solar array utilizes available “Standard” triple junction GaAs solar cells (27.5% efficiency)
Presently evaluating four different Li-Ion cell technologies for applicability and selection
Deployment control for Solar Arrays and High Gain Antennas located in PSE
Autonomous deployment control when receiving 3/3 LV separation signals
Reaction Wheel power feed control autonomously activated with 3/3 LV separation signals
Timer/spacecraft computer control backup when 2/3 separation signals received
Inter-component harnessing provided by Electrical Systems

5. SDO EPS Preliminary Design
Solar Array Module
Power Circuits (10 per wing)
Observatory Loads
Battery Module
Battery Enable Relays (2)
Deployment (S/A and HGAs) and Prop Pyro Enable Bus
Power Subsystem Electronics
Output Modules
Thermistors
Cell voltages
SDN
Solar Array
Thermistors
Voc Sensor
Isc Sensor
Coarse Sun Sensor
Deployment Module
Li-Ion Battery
Redundant 1553B Communications

6. Significant EPS Changes Since SCR
Solar Array sizing increased to accommodate added load and margin requirement (increased to 6.22 sq. meter solar cell area [7.7 sq. meter total array area])
Added Pyro Valve power bus enable FET for Propulsion System
Increased baseline battery capacity to 100 amp-hour to accommodate increased load demand
Replaced Subsystem Power Node (SPN) with Power Converter Card (PCC)
Incorporated cPCI backplane for PSE
Thermistor data processing from Instruments and Thermal Subsystem
Incorporation of observatory commanded relay to open battery circuit on orbit (over-discharge condition)

7. EPS Documentation Status
Power Subsystem Requirements Specification – Under project configuration control
GSE/EGSE Requirements – Draft completed, in preliminary review
Power Subsystem Development Plan – Baselined, signature cycle complete
Solar Array SOW/Specification – In Process (Completion by RFP Release)
Solar Array Qual/Acceptance Plan – In Process (Completion by RFP Release)
Solar Array Mechanical/Electrical ICD – In Process (Completion schedule after vendor selection)
PSE Requirements Specification – Submitted to CM, in project review cycle
PSE HW/SW ICD – Draft completed, in preliminary review
PSE Flight Software Requirements – Draft completed, in preliminary review
Battery SOW/Specification/Qual Acceptance Plan – In Process (Completion by RFP Release)
Battery Mechanical/Electrical ICD – Draft (Completion scheduled after vendor selection)

8. Component Preliminary Allocations
PSE
Mass
PSE CBE = 42kg (44.8kg Allocation)
Power
PSE Electronics CBE = 52W (54W Allocation)
Efficiency Losses = Sunlight 38W/ Eclipse 7W (Allocation 43W/9W)
Battery
Battery CBE = 42.5kg (45.3kg Allocation)
N/A
Solar Array
Solar Array (cell related only) CBE = 12kg (12.8kg Allocation)
Diode Losses CBE = 20W (22W Allocation)
EPS Subsystem
Power Margin 9.18% (Sunlight)/6.24% (Eclipse)
Mass Margin 6.74%

9. Performance Analysis Conditions
EPS Performance evaluated based on “nominal” and “worst case” conditions
Nominal calculations based on EOL conditions with no failures
Failure cases identified and examined:
Loss of one solar array circuit (~5% loss of array power)
Loss of one digital shunt circuit (~5% loss of array power)
Loss of one PSE PWM circuit (~10% loss of array power)
Loss of one battery cell (~12.5% loss of array power)*
Worst case performance analysis uses loss of large battery cell as benchmark for analysis
Battery vendor selection will heavily impact worst case analysis
Use of small cells will effectively eliminate a failed battery cell as worst case
* Loss of battery cell case assumes large Li-Ion cells selected, otherwise worst case is loss of one PWM circuit

10. Worst Case Energy Balance – EOL
1120 Watt Spacecraft Eclipse Load

11. EPS Performance Summary
Vast Majority of SDO mission life will be operation in full sunlight (319 days/year)
Battery voltage clamped at commanded value during sunlit period between eclipse seasons
Two orbital eclipse seasons each lasting 23 days
Maximum Eclipse of 72 minutes
Preliminary Sizing of EPS components provide sufficient design margin to meet SDO mission needs
Worst Case Analysis indicates EOL output of array over 1450 watts
Worst case battery DOD calculated at 51.2%
Battery performance characteristics will vary between available vendors
Performance calculations based on “typical” performance of Li-Ion battery
Selection of of small Li-Ion cell battery design will eliminate battery cell failure as worst case

12. EPS Redundancy Approach – Overview
Power Subsystem Electronics (PSE)
Redundant PSE bus architecture (including redundant processors)
Redundant Output Modules
Redundant Power Converter Cards
Redundant Deployment Modules
Solar Array Modules implement dual control and low voltage power interfaces
Redundant PWM (1/2) fine control shunts
Digital shunt redundancy allows operation on 15/16 shunts
Double insulated bus (some exceptions as noted and tracked in FMEA)
Hardwire command capability for critical switch functions
Solar Array (SA)
Designed to operate on 18/20 solar array circuits (loss of PWM circuit)
Diode isolated strings
Redundant Voc and Isc sensor circuits
Redundant PRTs on both array panels
Li-Ion Battery (BAT)
Designed to operate on 7/8 battery cells
Redundant PRTs on battery

13. PSE Design – Power Bus
One Power Bus – Single Fault Tolerant

14. EPS Design – Heritage
SDO PSE Based on MAP / EO-1 Power System Electronics
Direct Energy Transfer System - Battery dominated bus
SA Shunt with PWM Boost Converter for fine control
Re-use of much of Solar Array Module circuits
Re-use of most of Output Module design and components
Re-use of most of Battery Module design and components
New for SDO Mission
New Deployment Function
Higher Power (approximately 2X MAP power requirement)
Redundancy / Single Fault Tolerance
Each SAM must be internally redundant, can only lose 2/20 circuits
Circuitry to switch bus/software control over from one processor to the other
Mission Life
5yrs
Radiation
High radiation environment
Some parts substitution on re-used circuits
Leveraged Design Heritage
Solar Array Cells (Multiple Programs)
Battery *** (AEA - PROBA, Mars Express, XTRV, BEAGLE, Lithion – MER)
*** Battery vendor not selected, missions listed are from various potential vendors

15. LiIon Battery Risk/Mitigation (Risk #36)
Lithium Ion Battery (New Technology): Application usage of LiIon battery technology is still relatively new for space applications. Without significant test data to base heritage usage, concern (risk) that battery technology is not capable of meeting mission requirements.
Have undertaken extensive test program utilizing cells from four major LiIon battery vendors (with space cell applications) to determine simulated mission operation performance/degradation characteristics. Testing based on actual simulated GEO test profile (72% DOD, twice yearly 42 day eclipse season). Test battery program to be continued on selected vendor test battery and potential back up vendor until SDO launch (will have nearly 5 years on all tests by 4/08).
Detailed Test Battery Mitigation Plan
Perform simulated GEO testing on representative battery test packs to evaluate chemistry performance
Temperature: 20+/- 10 degrees C
Eclipse Period
Discharge: 0.6 C rate, for shadow period (maximum eclipse is 72 minutes)
Charge: C/2 rate, for remaining portion of the day, with TBD voltage clamp
42 days in length
Solstice Period
140 days, pack voltage maintained at TBD voltage clamp (70% SOC)
Prior to each eclipse season, the battery is charged up to 100% SOC using C/20 charge rate, to specified voltage clamp pertaining to vendor recommendation
Test program intended to retire risk completely by launch

16. Candidate Test Bed LiIon Life Testing: GEO
AEA 80 Ah Test Battery
JSB 100 Ah Test Battery
Saft 80 Ah Test Battery
Yardney 100 Ah Test Battery

17. EPS Reviews to Date
EPS Preliminary Design Review (PDR) – December 19, Completed
PSE Peer Review – December 16, Completed
SDN BB Peer Review – December 15, Completed
PCC/LPSC Peer Review – December 8, Completed
PCC/LPSC Requirements Review – October 8, Completed
SDO Spacecraft Concept Review (SCR) – April 8-11, 2003 – Completed
RFA Status
Total of 27 RFAs written/submitted at EPS PDR
Of the 27 RFAs Submitted
18 RFA Responses Submitted for Closure
9 RFA Closure Plans Developed/Submitted
No Significant Issues Identified
Other Action Item Status
All other EPS Related Action Items Either Closed or Recommended for Closure (Closure Plan/Response Submitted)

18. EPS Component Verification/Development Testing Planned
PSE (ETU) (In-house Design/Fab)
ETU/Flight card level safe to mate
ETU/Flight electrical performance verification at box level
ETU Electrical performance testing over temperature at box level
ETU Limited EMI/EMC Testing
Flight electrical performance testing/environmental testing (vibe, T/V, Full EMI/EMC testing)
Battery (Vendor Supplied)
Test battery performance (simulated GEO orbit)
Test/Qual/Flight cell selection/matching based on performance and characterization testing
Test/Qual/Flight battery electrical performance testing over temperature
Qual/Flight battery electrical performance testing/environmental testing (vibe, T/V, EMI/EMC testing)
Solar Array (In-House Substrate Fab/Vendor Supplied Solar Cells & Wiring)
Qualification panel development and testing (fab & repair procedures, electrical, T/V & thermal cycling)
Flight array electrical performance verification at vendor (flash testing)
Flight array environmental Testing (Vibe, T/V bakeout)
Flight pre/post environmental electrical performance flash testing
Flight electrical performance verification at GSFC to include “Hot Flash” testing
Periodic illumination lamp testing

19. SDO Power Subsystem Schedule
Subsystem & Element
CY 2003
CY 2004
CY 2005
CY 2006
CY 2007
CY 2008 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4 Q1 Q2 Q3 Q4
4/8
8/03
3/04
5/04
2/05
MISSION MILESTONES
LAUNCH
SRR/
SCR
ICR
PDR CR
CDR
PER
PSR
Power
12/19/03
12/04
PDR
CDR
Power Subsystem Electronics (PSE)
1 = Spacecraft Integration
= Schedule Reserve
2 = Instrument Integration
3 = Environmental Testing
4 = Launch Site Operations
Design
ETU PWB Buy
ETU Assy
ETU Test
FLT Fab & Assy
FLT Box Test
Batteries
Spec/SoW
Award to Vendor
Qual Batt Fab Assy Test
Flt Batt Fab & Assy Test
Solar Array
Spec/SoW
Procure
Qual Panel Design
Qual Panel Procure
Qual Panel Fab & Assy
Qual Panel Test
Flt Fab, Assy & Test
Spacecraft I&T 1 2 3 4
Launch

20. Issues/Conclusion Issues/Concerns Conclusion
No EPS major issues/concerns have been identified
Conclusion
SDO Electrical Power Subsystem Preliminary Design is complete
EPS meets all imposed requirements with adequate margin and is ready to proceed to the Critical Design phase

21. Backup Material

22. EPS GSE/EGSE GSE/EGSE arrangement for SDO EPS Test/Operation
Solar Array Electrical Simulator (SAES) (Terra) - Electronics Loads (std. lab supplies) - A/C Cooling Unit
Battery Simulator (XTE/TRMM) - Power Supply (std. lab supplies)
Assist Workstation (from S/W) - Battery GSE (BGSE) (new)
Power Subsystem Electronics (PSE)
Solar Array Electrical Simulator
(SAES)
20 Solar Array Circuits (maximum 3.5A/circuit)
Battery GSE
(BGSE)
Battery Cell monitoring and reconditioning
Li-Ion Battery (BAT)
Assist Workstation
Electronic Loads
Battery Simulator (BSIM)
(BSIM will only be used when battery not available)
Power
Telemetry/
Command
Power Supply
Battery Enable Relay (2 in parallel)
Startup Power
Component/lab use
A/C Cooling System
Solar Array Wing 1 & 2

23. PSE Design – Redundant cPCI Backplane

24. End of Life Losses Off Angle of 4 Degrees, -.2% Cell Mismatch, -1%
Coverglass Assembly, -2.6%
Thermal Cycling, -1%
Hard Particle Radiation, ~ -10%
UV, -1%
Debris & Micrometeorite, 0%
Measurement Error, -2%
Contamination, 0%

25. 71.2C SDO Low VArray EOL Power
16.5% (Cells Powering S/C)
Peak Power 33.1V
26.1V
26.1V (Controlling Case) V
29.6V
35.2V
35.2V

26. -170C Hi VArray EOL Power Peak Power 2447W @ 55.7V 47.1A @ 26.1V
26.1V ok, as array heats quickly V
29.6V ok, as array heats quickly
35.2V
35.2V