|ELECTRICAL POWER SYSTEM
The SNOE Electrical Power System (EPS) is a direct energy transfer system
that generates energy, stores energy for use during orbit eclipse, and
controls the distribution of energy to the required S/C and payload systems.
Design goals for the EPS include generation and distribution of reliable
28V (plus or minus 4V) power to the S/C and payload with a greater than
20% end of life (EOL) margin. Much of the SNOE power system design is modeled
after the Solar Mesosphere Explorer (SME) spacecraft EPS.
Power is generated by body mounted panels of solar arrays. There are twelve rectangular panels, each containing 2
array circuits (strings) for total of 24 strings. The strings consist of
seventy-four 2.3 cm x 4.1 cm silicon photovoltaic cells. Diode isolation
of the individual array circuits from the bus is used to prevent a failure
in an array string from causing a failure of the entire power generation
system. Assuming one-year degradation of 10% and 30 degree solar beta angle,
the solar arrays will generate
greater than 42 watts orbit average power at EOL.
The solar cells were purchased from Heliokinetics of Palm Springs, CA.
The 10 Ohm-cm, 8 mil. thick cells are AR coated and delivered with pre-installed
cover glass and interconnection butterfly fittings. Criteria and procedures
for selection of the individual flight solar cells include testing for
cell open circuit voltage and short circuit current. Assembly of the arrays was performed by
student assemblers (trained by industry consultants) at LASP under the
supervision of LASP QA personnel.
Energy required for peak loads and during the eclipse portion of the
orbit is stored in two 22-cell 4-Amp-hour Nickel Cadmium (NiCd) battery
packs. The batteries use Sanyo D cells, tested, integrated and qualified
by a flight proven Ball Aerospace process. With exception of telemetry
transmission, the S/C and payload have static power usage profiles. There
are no diurnal or seasonal experiment scenarios. The transmitter draws
28W during transmission, which occurs twice per day for less than 20 minutes
per pass. Total orbit average power required is 35 watts. This power usage
produces a battery depth of discharge
(DOD) of less than 10%. Typical life for a NiCd battery operated at <20%
DOD is >20,000 charge/discharge cycles (>3 years).
A shunt regulator is used to prevent overcharging of the battery and
the associated generation of excess heat. Shunt regulation is performed
by clamping the bus voltage to a level set by predetermined voltage and
temperature curves (V-T control). Bus clamping action (regulation) is implemented
by a two mode shunt controller which uses a combination of conventional
linear shunt regulation (fine control mode) along with solar array string
circuit switching (coarse control mode). The coarse control is achieved
using switched array string circuits controlled by the V-T controller in
a closed loop fashion. There are three switching circuits of six strings
each, with the six strings distributed one to each S/C face to prevent
current fluctuation when one or more circuits are switched off. Fine control
is performed by a linear regulator that shunts excess power into resistors
mounted on S/C surfaces. The V-T controller uses 4 temperature voltage
curves selected by uplink command. Power distribution includes two primary
power busses; the essential bus and the non-essential bus. Circuits on
the essential bus are the EPS itself, the command receiver/demodulator,
and the S/C processor. All other loads are considered non-essential and
are therefore disconnected when the bus voltage monitor detects a low voltage
condition. Because EPS reliability is of primary importance to the spacecraft,
the power control unit has some redundant
components in critical areas.