MAXIM Power Subsystem Diane Yun Vickie Moran NASA/GSFC Code (IMDC) 8/19/99

Mission Description  MAXIM is a straight-forward mission for power system design.  The only area of concern is the uncertainty in the instrument loads; however, the mission is not weight, surface area, or launch vehicle volume constrained which means that larger solar arrays or batteries could be incorporated later if needed.  The MAXIM mission consists of two spacecraft formation flying in a fly- away orbit which is essentially heliocentric resulting in no eclipse periods.  Each spacecraft will be maintained solar inertial (i.e. having one axis pointed at the sun continuously).  The “Optical” spacecraft contains the optical mirrors.  The “Detector” spacecraft maneuvers to maintain its detector in line with the focus of the mirrors on the “Optical” spacecraft and ~450km from the “Optical” spacecraft.

Loads Rack-Up & Assumptions  Optical Spacecraft--521W including 20% contingency ë259W Instrument Load ë262W Spacecraft Load  Detector Spacecraft--566W including 20% contingency ë313W Instrument Load ë253W Spacecraft Load  Both spacecraft have identical attitude control sensors and actuators including: CSS, IRU, 4 RWs, Thrusters, and a Star Tracker.  The reaction wheels are assumed to operate at an average of 18W each except when maneuvering for new targets when the power goes up to 80W each wheel (approximated for power system sizing as 2 hours each day).  The Power System Electronics (PSE) is a Direct Energy Transfer (DET) architecture similar to MAP which is at L2. Power Distribution is included in the MAP PSE.  The detector spacecraft has a low power S Band Receiver and Transmitter (20W). Both are on continuously.  The optical spacecraft has a low power S Band Receiver and Transmitter (20W). The receiver (5W) is on continuously. The transmitter is only on for maneuvers to new targets (approximated for power system sizing as 2 hours each day).  The optical spacecraft has a X Band Receiver and Transmitter. The receiver (5W) is on continuously. The transmitter (32W) is on ~ 1 hour each day.  The C&DH is estimated to consume equivalent power (30W) on each spacecraft and incorporate all ACE functions.  We assume less heater power for the detector spacecraft (15W) than the optical spacecraft (30W) because of the lower energy density.

Loads Rack-Up

Power System Sizing Philosophy  The system was sized to provide energy balance on a 24 hour basis.  There are no eclipses but we still need to carry a battery to supply power in the launch vehicle until spacecraft sun acquisition on orbit. We minimize the solar array area required by using the battery to supply supply transient peaks in the load over the 24 hours.  The solar array is sized to provide the average load and recharge the battery over the 24 hour period.

Battery Sizing  The peaks identified to date (for the Optics spacecraft) are: ëThe optics motors during a mirror adjust. (6 motors running simultaneously at 10W each 2 hours. SA is sized for 5W average; energy out of battery=(60W-5W)*2hrs=110Wh) ëThe reaction wheels slewing the spacecraft for target acquisition. (4 wheels running at 80W each for 2 hours. SA is sized for 92.7W average; energy out of battery=(320W-92.7W)*2hrs=454.6Wh) ëThe X Band Transmitter running at 32W for 1 hour each day. SA is sized for 1.3W average; energy out of battery=(32W-1.3W)*1hr=30.7Wh ëThe S Band Transmitter on the Optics S/C (15W) for target acquisition 2 hours each day. SA is sized for 1.25W average; energy out of battery=(15W-1.25W)*2hrs=27.5Wh ëTotal Energy Out Of Battery Each Day=622.8Wh=22.2Ah ë~60Ah battery provides a maximum DoD of 46% for 3 year life % DoD max is recommended for GEO (24 hour charge/discharge cycles).  The peaks identified to date (for the Detector spacecraft) are: ëThe reaction wheels slewing the spacecraft for target acquisition. (4 wheels running at 80W each for 2 hours. SA is sized for 92.7W average; energy out of battery=(320W-92.7W)*2hrs=454.6Wh) ëTotal Energy Out Of Battery Each Day=454.6Wh=16.2Ah ëBecause the energy out of the battery could be much higher than calculated above due to detector alignment with the optics spacecraft and because there is no weight issue with the launch vehicle, we have specified the same battery size for the detector spacecraft as the optics spacecraft. Further analysis of the power required for the detector alignment maneuvers needs to be done.  10 years from now it is reasonable to assume that rechargeable Lithium Ion batteries will be available in large capacity sizes with energy densities of ~120Wh/kg; Battery Weight: 14kg

Solar Array Sizing--Optical Spacecraft  The solar array is body-mounted on the half of the cylindrical surface that faces the sun.  The solar cells are mounted to a band around the body of the cylindrical spacecraft. The band dimensions are 3.8m dia x 2.0m length. The projected area is 7.6m 2.  We assume that 85% of that area is available for mounting active solar cells.  Sizing Results (assuming 6% loss in efficiency from BOL to EOL): ëTechnologyProjected AreaSurface AreaWeight * ROM Cost ëSi6.76m m kg1038K ëGaAs4.33m m kg2224K ëTriple-Junction GaAs3.33m m 2 8.9kg2231K ë*Does Not Include Substrate (included in mechanical budget)

Solar Array Sizing--Detector Spacecraft  Because the body of the spacecraft is too small to support a body-mounted solar array, the solar array is mounted to fixed, deployed panels. The panels are maintained normal to the sunline by the spacecraft.  The projected area is 6m 2.  We assume that 85% of that area is available for mounting active solar cells. ëTechnologyProjected AreaWeight*ROM Cost ëSi7.16m kg700K ëGaAs4.58m kg1499K ëTriple-Junction GaAs3.53m kg1503K ë*Includes Substrate

Solar Array Output--Body Cells  The solar array was sized assuming 11.1% loss in power from BOL to EOL & 90°C SA temperatures. ëWe do not have the capability to perform radiation analyses for non-Earth orbiting spacecraft. Since weight is not a problem a thicker coverglass can be applied to the cell to protect against radiation damage.