IEEE Geoscience and Remote Sensing Magazine - June 2013 - 58

Research Institute's flight-proven, well-characterized, thermal design techniques. The thermal control design provides thermal stability and minimizes thermal gradients
through an integrated design of multilayer insulation blankets (MLI), surface treatments, and localized radiators. The
arrangement of internal equipment is used to aid thermal
control and eliminate the need for supplemental heaters
except for Standby/Safe Hold operations.
Results from our thermal analysis were used to size the
thermal radiators (EOL). The primary radiator is located on
zenith surface in the S/A gap along the Observatory center
line, with a second radiator on the nadir baseplate. These
locations are chosen to provide a direct, cohesive thermal
conductive path to the primary observatory dissipative
loads. The radiators are coated with 5 mil ITO/Tef/Ag, while
MLI is used on non-radiating external surfaces.
4. avIonIcS
The CYGNSS Avionics consists of four boards, portions
of the EPS and CDS. The boards include the Peak Power
Tracker, the Low Voltage Power Supply, the Centaur single
board computer, and the Flexible Communication Platform radio. A block diagram of the avionics unit is shown in
Figure 9. The avionics unit does not include a box; instead,
the microsat structure itself provides mechanical mounting
and electrical interconnects over a backplane and cables.
4.1. elecTrical PoWer subsysTeM
The EPS design performance provides robust margins on
all requirements. The EPS is designed to perform battery
charging without interrupting science data acquisition.
Solar array
The EPS is based on a 28!4 Vdc primary power bus with
electrical power generated by an 8-panel rigid solar array
(S/A). The S/A design is composed of solar panels, hinges,
and deployment actuators. Four of the eight panels are
"z-folded" for launch. Flight-qualified, triple-junction solar
cells are arranged with an 84% packing density on the solar
panel substrates, including cover glass to improve their thermal performance and ground handling robustness. The
0.71 m2 total area S/A provides a 30.3% margin during max
eclipse periods (35.8 min). Full mission duration simulations were performed to analyze worse case solar Beta cases
(58). The design provides 43.4% margin during these periods. When stowed, the z-fold design of the S/A allows the
solar cells to face outward, combining with the two supplemental ram/wake S/As to power the microsat indefinitely in
Standby mode before S/A deployment (22% margin).
BatterieS
Electrical power storage for eclipse operations is provided
by two 1.5 A-hr Li-ion 8s1p batteries connected directly
to the primary power bus. The batteries are configured
for 3 A-hr (EOL) at 28.8 Vdc nominal. Temperature sensors, and bypass diodes (to withstand a failed cell) are
58

included in the battery assembly. Battery performance
models were used to analyze the CYGNSS mission with
predicted EOL nominal battery state-of-charge being
87.6%. Battery charging uses a constant current, voltagetemperature limited charge scheme based on four stored
profiles matched to the CYGNSS battery. Charging is also
Coulomb limited to 120% of discharge level. The primary
power bus voltage is modulated to maintain charge current and termination voltage. The Coulombic charge limit
is tracked with an A-min integrator and when the level
exceeds 1.2 # IdisTeclipse (Amin), battery charging levels are
reduced to C/100.
Peak Power tracker
Battery charge regulation for the CYGNSS EPS is a peak
power tracking (PPT) type regulator. The PPT board,
developed using SwRI internal funds, matches S/A conductance to the Observatory load through pulse-width
modulation (PWM) using an optimization control circuit
that integrates S/A W-sec over a preset period of time. The
PPT includes a ground support equipment (GSE) interface that serves as the connection point for ground power
and battery maintenance, conditioning, and pre-launch
trickle charging.
The PPT unit is based on a 40W DC-DC converter,
which produces 28!4 Vdc from a solar array voltage of 36 to
72 Vdc. The design was produced with multiple missions in
mind, from a long-duration, intense radiation environment
to a short, LEO mission, CYGNSS being toward the latter of
these two extremes. The DC-DC converter output voltage is
modulated by the PPT and battery charge regulator to meet
load power and battery charging demands. Power from
the solar array flows into the PPT through an over current
protection fuse, current sense resistor and EMI filter. S/A
current and voltage are sensed and conditioned before connection to an analog multiplier within the PPT circuit. The
analog multiplier converts these signals into instantaneous
S/A power, which is processed by the PPT watt-second integrator to track the power peak. The PPT circuit generates an
error signal (PPT Error), which is used to provide supervisory control of the DC-DC converter in conjunction with
the battery charge regulator.
Housekeeping power is provided by a high input voltage
linear regulator, which provides +16 Vdc for control circuit
power and midpoint bias of +8 Vdc to operate single supply
operational amplifiers.
Battery charge regulation consists of programmable
charge current and end-of-charge voltage settings, which
are each controlled via opto-isolated 4-bit interfaces.
The opto-isolators are set up for 3.3 Vdc CMOS drive
levels from the Centaur interface. No flight software is
required for the control electronics, except for configuration control.
The PPT is also used to switch +28 Vdc bus voltage to
spacecraft components, including the S/A deployment actuators, the DMMI, heaters, and momentum wheels.
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Table of Contents for the Digital Edition of IEEE Geoscience and Remote Sensing Magazine - June 2013