DINO PDR 1132020 DINO Electrical Power Subsystem PDR

DINO PDR 11/3/2020 DINO Electrical Power Subsystem PDR Briefing Mike Wong Jennifer Michels Kevin Mc. Williams July 29, 2003

DINO PDR 11/3/2020 Spacecraft Level Requirements • • • Placed on-orbit from Internal Cargo Unit aboard Space Shuttle LEO, International Space Station orbit (51. 6º, ~400 km altitude) Mission duration between 6 and 12 months 55 minute sun-side, 35 minute eclipse 10 W power during safe mode DINO, Tether, FITS, and Aero-Fins deployed

DINO PDR 11/3/2020 Power Subsystem Requirements • • • The system will weigh < 2. 25 kg The system will communicate with C&DH through a shared RS-232 line The system will provide varying current levels to ADCS (at least 3) The system will provide a 28 V line to structures for deployment mechanisms The system will provide C&DH and COMM receiver continuous power The battery cells will last for 6, 000 day/night cycles The system will provide 30 W nominal power during the day, 15 W at night EPS will provide 5 V, +/-12 V and 28 V lines The system will receive instructions from C&DH The system will provide data to C&DH for processing The system will operate at temperatures between 0 and 40 degrees Celsius with a nominal operating temperature of 20 degrees Celsius

DINO PDR 11/3/2020 Key Requirements • The satellite will be unpowered while in the shuttle payload bay • All inhibits will verify their current status • Current will be monitored on primary inputs and outputs of the EPS board • The battery will have 3 independent inhibits; one on the negative lead and two on the positive lead • A fuse will be located in line with the ground leg of the battery • Voltages will be monitored on the busses, battery, and solar arrays • Solar arrays will be isolated using 1 independent inhibit • The spacecraft loads will be isolated using 1 independent inhibit • The EPS will interface with the Ground Support Equipment (GSE) to perform test operations • All components will be de-rated for worst case conditions

DINO PDR 11/3/2020 Safety Requirements - Battery • Battery; JSC 20793 – Manned Space Vehicle Battery Safety Handbook • Lithium battery systems must be two-fault tolerant – requires 3 independent inhibits • Stack packaging must protect against corrosion and shorts due to electrolyte leakage • Cell vents cannot be oriented downward at any time during launch on the shuttle • Cells must be provided with individual temperature monitoring capability and shunt diodes • Charging/discharging must be done in such a way as to not damage the cells • The cells will be maintained above 3. 2 V cell voltage to prevent current reversal • The stack will be nominally charged to 95% of capacity to prevent overcharge of individual cells • Battery assembly must be maintained between 0 and 40 degrees Celsius

DINO PDR 11/3/2020 Safety Requirements - Other • Wiring • All wires/fuses must be de-rated for worst-case conditions • Wire harness must not interfere with or be damaged by structure • Wire bundles must not overheat • All inhibits must be able to have status verified from GSE without spacecraft power

DINO PDR 11/3/2020 Electrical Power System Block Diagram SCIENCE Cam 1 COMM HOP SW Radio 2 Cam 2 Radio 1 Imaging GSE 15 V Prim SA Back SA Wireless TNC TR 1 RS-232 I/O S Control Card 5 V TR 3 12 V EMC 1 EP 5 V TR 2 Mag. STRUCTURES HOP C&DH ADCS Sun Sen. Battery EMC 2 28 V EMC 3 EMC 4 Boom

DINO PDR 11/3/2020 Inhibits Layout Diagram EPS Board • Main EPS uses separation switches from ICU to trigger system start-up • Connect to two FETs on the high leg of the battery, one on the ground leg • Method to get power into EPS with un-charged battery is via a dedicated SA string • Solar Array (SA) string switches will act as SA inhibits • Subsystem load switches will act as spacecraft load inhibits

DINO PDR 11/3/2020 Electrical Power System Description • Direct energy transfer system • Dual Solar-Array technologies (2 solar arrays) • • Conventional body-mount crystalline cells Thin-Film (FITS) Flexible Array • System will switch solar array strings and spacecraft loads as instructed by C&DH • • Will require 4 switches for the FITS Will require 16 (TBR) switches for the backup SA 4 switches capable of transferring less than 2 Amps 12 switches capable of transferring less than 1 Amps • • • System will use 4 A-hr Lithium Polymer cells at 3. 7 Volts for power storage 4 Cells will be in series, for a 15 V +/- 2 V battery assembly Battery charging will be accomplished using several charging modes: 1. Nominal sub-capacity charging; entire stack is charged to 95% capacity 1. This will be the default mode 2. Full-charge; entire stack charged to 100% capacity 3. Cell conditioning over-charge; as stack ages, it will lose capacity, which can be restored by judicious over-charging 4. Temperature and voltage monitoring will be the primary means of controlling cell charging

DINO PDR 11/3/2020 Switch List Function Number of Switches Amps Volts ADCS 3 0. 500 5 COMM RX 1 1. 500 12 Thermal 3 (TBD) Cam-1 1 0. 200 5 Cam-2 1 0. 200 5 HOPS 1 0. 643 28 Memory Hinges 4 0. 357 28 Patch Antenna 1 (TBD) COMM TX 1 0. 100 5

DINO PDR 11/3/2020 Electrical Power System Mission Modes • Start-up Mode • • • ADCS, C&DH, COMM on Activate High-Output Paraffin actuator (HOPS) Deploy tether • tip-mass experiment views other deployments • • Deploy FITS, primary power source Deploy aero fins by activating Electric Memory Composite (EMC) hinges Normal Mode • All systems un-powered Inhibits must trigger start-up event EPS activates, followed by C&DH, and COMM Charge battery immediately Deployment Mode • • • 30 W power, two sub-modes • Transmit Mode, cameras off, COMM in transmit (TX) and receive (RX) • Science Mode, cameras on, COMM in RX • Night Mode • • • C&DH, COMM on 15 W nominal power provided to the spacecraft from the battery Safe Mode • All non-essential systems off, C&DH, COMM in RX and EPS on

DINO PDR 11/3/2020 DINO Power Consumption During a Regular Cycle Eclipse Begins NOTE: Plot begins and ends as DINO exits eclipse

DINO PDR 11/3/2020 DINO Available Power over 1 Year Power Output, Watts Total Wattage Produced by Solar Panels for 4 Sun Positions in ISS Orbit Nominal max. Power Consumed Time, Minutes

DINO PDR 11/3/2020 Electrical Power System Control Card Block Diagram C&DH (RS-232) Cmd/Monitor line EPS Control Card Power line 5 V Reg FITS SA Interface Charge Control & Monitoring +/- 12 V Reg 28 V Reg Body Mount Battery Subsystem Interface PMAD • Provides SA switch interface • Provide charge control/Monitor • Provides C&DH RS 232 interface • Provides secondary Power • 5 V • +/-12 V • 28 V • Provides 16 load switches to the spacecraft

DINO PDR 11/3/2020 Ground Support Equipment Requirements • GSE will exercise all inputs and outputs of the EPS controller card • GSE will duplicate all EPS-related functions of the flight computer • GSE will be able to power up system while all system inhibits are in place • GSE will provide external power to EPS controller card • GSE will verify SA and S/C load switch status • Battery will be charged and discharged by GSE • GSE will monitor status of all system inhibits • GSE will read data from sensors • Voltage sensors • Current sensors • Thermistors Credit: http: //sohowww. nascom. nasa. gov/

DINO PDR 11/3/2020 Battery Schematic Credit: Valence Electronics www. valence. com Battery Lithium-Polymer cells Prismatic design 4 A-hr capacity 4 cells in series 4 flight temp monitors 4 GSE temp monitors Cell voltage to GSE Battery voltage to flight

DINO PDR 11/3/2020 Solar Array CIGS (what ‘s this mean? ) Cell Interconnects 5 Bonded CIGS PV on Stainless Steel Substrate Etched Structural Bond Area Masked Electrical Bond Area Electrically Conductive Adhesive High Strength Space Qualified Dielectric Adhesive

DINO PDR 11/3/2020 Blanket Configuration Y (8 Cells) X 1 Cells String 6 Modules interconnected by MSI using reinforced Kapton tape to form a string of 48 cells. 2 strings per solar array wing. Array-Wing 2 Strings mechanically joined by MSI using reinforced Kapton to form a wing – 0. 552 m 2

DINO PDR 11/3/2020 Solar Array Electrical Properties Beginning of Life End of Life 25° Celsius 78° Celsius – Radiation Degradation 1. 10 Voltage – String 17. 4 Volts 13. 0 Volts Current – String 1. 4 Amps 1. 27 Amps Power – Total Watts 85 Watts 64 Watts 5. 70% 4. 60% Area Array Efficiency

DINO PDR 11/3/2020 Command & Monitors List EPS Mode Commands Subsystem commands Start-up Deployment Day-side Normal Ops Science Transmit Night-side Normal Ops Safe Mode ADCS levels COMM: RX/TX Science on/off C&DH: send data, reset Structures: power to HOPS or hinges SA’s: turn strings on/off Spacecraft Monitors Thermistor on each cell in battery assembly Thermistor on EPS card Voltage and Current sensors for Battery Assembly feed One string on each SA Each sub-system Voltage for each Cell in stack Inhibit/Switch status

DINO PDR 11/3/2020 Test Plan • • Battery to be certified by either Ball Aerospace & Technologies Corp. (BATC) or cell vendor Circuitry developed by sub-contractor Thermal-vacuum and vibration testing will be carried out by CSGC students using BATC facilities GSE developed by CSGC • • • Verify system inhibits Verify voltage min/max Provide SA simulation Verify battery charge monitors Verify commands/monitors Perform step load testing Monitor temperatures Environmental testing Interface with monitor Exercise all inputs, outputs, and logic operations of EPS Credit: NASA/JPL http: //mars. jpl. nasa. gov/mer/

DINO PDR 11/3/2020 Parts Source List • • • Batteries – Provided by BATC Wiring Harness – built by CSGC EPS Card – Provided by BATC and sub-contractor Fuses – Provided by CSGC Connectors – Available everywhere GSE – Built by CCSG Credit: NASA/JPL/Cornell/ASU http: //athena. cornell. edu/the_mission/ins_minites. html Credit: Broad Reach Engineering http: //www. broadreachengineering. com/eps. html

DINO PDR 11/3/2020 Trade Studies Study Status Battery chemistry Ni. Cd/ Ni-MH vs. Lithium Ion-Polymer Completed Tether power source Independent source vs. Power line from main satellite Completed 28 volt bus Boost converter vs. Thermal battery In work

DINO PDR 11/3/2020 Issues and Concerns • • • Certification of battery cells for spaceflight Starting up the spacecraft when leaving the shuttle (initial power-up) Monitoring inhibits and sensors via GSE with no internal power provided Obtaining a contract agreement with BATC Use of AFRL battery cells, if BATC can not help with battery certifying