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Miniature X-Ray Solar Spectrometer (MinXSS): A Science-Oriented, University 3U CubeSat

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The Miniature X-ray Solar Spectrometer (MinXSS) 3U CubeSat is scheduled to launch to the International Space Station (ISS) on ORB-4 in November 2015. MinXSS was designed and developed as part of the Aerospace Engineering Sciences (AES) graduate projects course at the University of Colorado in Boulder (CU) with significant facilities and professional support from the Laboratory for Atmospheric and Space Physics (LASP). Some of the spacecraft design is heritage from the highly successful Colorado Student Space Weather Experiment (CSSWE) CubeSat, which was the result of the same AES/LASP collaboration. The project course was initially supported by AES and NSF, and MinXSS flight has been funded as a science mission by the NASA Heliophysics Division. This paper will provide an overview of the MinXSS science of observing the highly energetic solar soft X-ray (SXR) radiation, the educational program of involving students in design, building, and testing a CubeSat, and some lessons learned.
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Mason 1 29th Annual AIAA/USU
Conference on Small Satellites
SSC15-III-6
Miniature X-Ray Solar Spectrometer (MinXSS)
A Science-Oriented, University 3U CubeSat
James P. Mason, Thomas N. Woods, Gregg Allison, Matthew L. Cirbo, Seth Folley, Andrew Jones, Rick Kohnert,
Xinlin Li, Christopher Moore
Laboratory for Atmospheric and Space Physics, University of Colorado Boulder
3665 Discovery Drive, Boulder, CO 80303; 209-221-3766
james.mason@lasp.colorado.edu
Amir Caspi
Southwest Research Institute
1050 Walnut Street, #300, Boulder, CO 80302; 303-546-6351
amir@boulder.swri.edu
Scott Palo
University of Colorado Boulder
429 UCB, University of Colorado Boulder 80309; 303-492-4289
palo@colorado.edu
ABSTRACT
The Miniature X-ray Solar Spectrometer (MinXSS) 3U CubeSat is scheduled to launch to the International Space
Station (ISS) on ORB-4 in November 2015. MinXSS was designed and developed as part of the Aerospace
Engineering Sciences (AES) graduate projects course at the University of Colorado in Boulder (CU) with significant
facilities and professional support from the Laboratory for Atmospheric and Space Physics (LASP). Some of the
spacecraft design is heritage from the highly successful Colorado Student Space Weather Experiment (CSSWE)
CubeSat, which was the result of the same AES/LASP collaboration. The project course was initially supported by
AES and NSF, and MinXSS flight has been funded as a science mission by the NASA Heliophysics Division.
This paper will provide an overview of the MinXSS science of observing the highly energetic solar soft X-ray
(SXR) radiation, the educational program of involving students in design, building, and testing a CubeSat, and some
lessons learned.
MISSION OVERVIEW
The Miniature X-Ray Solar Spectrometer (MinXSS) is
a 3U CubeSat that began development as an aerospace
student project at the University of Colorado Boulder
(CU) and the Laboratory for Atmospheric and Space
Physics (LASP) in August 2011. The primary objective
of the science-oriented MinXSS CubeSat is to better
understand the energy distribution of solar soft X-ray
(SXR) emission and its impact on Earth’s ionosphere,
thermosphere, and mesosphere (ITM). With NSF
support in 2013 and subsequent NASA funding in
20142016, three MinXSS units have been fabricated
(Figure 1): a prototype and two flight models. The
prototype MinXSS has been valuable for early testing
and fit checks, and as extra unit for developing flight
software in parallel with other build activities. MinXSS
flight model 1 (FM-1) is ready for launch and is
manifested on the fourth International Space Station
(ISS) resupply mission by Orbital ATK (ORB-4), to be
launched on an Atlas V on 19 November 2015.
MinXSS FM-1 will be deployed from a NanoRacks
CubeSat Deployer on the ISS in January 2016, where it
will have an expected 512 month orbital lifetime,
dependent on atmospheric conditions. MinXSS FM-2 is
being planned for a higher altitude, longer mission in a
sun-synchronous polar orbit (SSPO) via a launch on the
Skybox Minotaur C launch in 2016. This section
provides an overview of the science objectives, the
history of the project, and the spacecraft subsystems.
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Figure 1: The MinXSS family (left to right): prototype unit, flight model 1, flight model 2.
Science Objectives - Overview
There is a rich history of solar SXR spectral
observations over the past three decades, but with a
significant gap of spectrally resolved measurements in
the 0.46 nm range.1 There were many new discoveries
about solar flares during the 1980s using solar SXR
spectral measurements from the DoD P78-1, NASA
Solar Maximum Mission (SMM), and JAXA Hinotori
satellites. For example, Doschek provides results about
flare temperatures, electron densities, and elemental
abundances for some flares during these missions.2 A
review of flare observations from Yohkoh and the
Compton Gamma Ray Observatory (CGRO), for the
hard (higher energy) X-ray (HXR) range, is provided
by Sterling et al.3 These earlier missions laid a solid
foundation for studies of flare physics and flare spectral
variability that the Reuven Ramaty High Energy Solar
Spectroscopic Imager (RHESSI) and the Solar
Dynamics Observatory (SDO) continue today for the
HXR and EUV ranges, respectively.4,5 With solar flare
spectral variability expected to peak near 2 nm, in a
range not currently observed by any spectrometer,
MinXSS measurements of the solar SXR irradiance will
provide a more complete understanding of flare
variability in conjunction with measurements from
RHESSI and SDO EUV Variability Experiment
(EVE).6,7
There are also nearly four decades of broadband (5-
10nm wide) SXR measurements that do not provide
spectrally resolved measurements. The very limited
spectral information from these broadband
measurements cannot quantify the specific spectral
energy distribution, nor directly quantify the varying
contributions of emission lines (bound-bound) amongst
the thermal radiative recombination (free-bound) and
thermal and non-thermal bremsstrahlung (free-free)
continua. These broadband measurements include,
among others, the two GOES X-Ray Sensor (XRS)
channels covering a combined band of 1.625 keV
(0.050.8 nm) and the even broader band of 0.212 keV
(0.17 nm) from several missions, including the
Yohkoh Soft X-ray Telescope (SXT, 19912001),
Student Nitric Oxide Experiment (SNOE, 19982002),
Thermosphere-Ionosphere-Mesosphere Energetics and
Dynamics (TIMED, 2002present), the Solar Radiation
and Climate Experiment (SORCE, 2003present), and
SDO (2010present).8,9,10,11,12 Broadband measurements
of solar SXRs have helped to resolve an outstanding
difference between ionospheric models and
measurements, such as the electron density from the
Haystack Observatory incoherent scatter radar at
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Millstone Hill. In particular, the SNOE solar
measurements were able to resolve the factor-of-4
difference between models and measurements because
the SNOE data indicated much more SXR irradiance
than had been previously thought.13 Additional
broadband SXR measurements have been made since
then; however, differences still remain in understanding
solar SXR spectral distribution and atmospheric
photoelectron flux. While smaller, these discrepancies
are still as large as a factor of 2 at some wavelengths, as
shown in Figure 2; the lack of spectral resolution in the
SXR range is thought to be the culprit for most of these
disagreements. For example, Peterson et al. show that
discrepancy between photoelectron measurements and
models were significantly improved with new EUV
spectral measurements down to 6 nm, and we anticipate
further improvement with new solar SXR spectral
measurements and atmospheric modeling with data
from MinXSS due to its ability to measure all
wavelengths in its spectral range simultaneously and
with the relatively high spectral resolution of 0.15 keV
FWHM.14
Science Objectives Solar Flare Studies
Spectral models of the solar irradiance (e.g., CHIANTI)
are needed in order to convert spectrally-integrated
broadband measurements into irradiance units.15,16
Detailed modeling to estimate the SXR spectrum during
a flare in April 2002 using a set of broadband
measurements from the TIMED Solar EUV Experiment
(SEE) was performed by Rodgers et al.6 The CHIANTI
spectral model is part of their analysis and is also
routinely used for processing these broadband
measurements. e.g., 17 While the CHIANTI spectra are
scaled to match the broadband SXR irradiance in data
processing, there are significant differences for
individual emissions lines between the CHIANTI
model and observations, often more than a factor of
two.18,19 Furthermore, there are concerns that CHIANTI
could be missing many of the very hot coronal
emissions lines, especially in the SXR range where
there are so few spectral measurements between 0.5 and
6 nm. Additionally, there are factor of 2 differences
when comparing the irradiance results from different
broadband instruments, which are worst during times of
higher solar activity (Figure 2). These discrepancies can
be partially explained by wavelength-dependent
instrument calibrations, but the greater contribution is
likely the lack of knowledge of how this dynamical part
of the solar spectrum changes as a function of
wavelength and time.
The MinXSS spectrometer, an Amptek X123-SDD,
flew on the SDO/EVE calibration rocket payload in
June 2012, and that measurement had a difference of
almost a factor of 8 below 2 nm as compared to the
CHIANTI model prediction based on SORCE XPS
broadband measurements.20 This rocket result was a
surprise considering that the SORCE-based CHIANTI
model prediction agreed with SDO/EVE measurements
down to 6 nm. Improvement of models of the solar
SXR spectra, which is only possible with calibrated
spectral measurements of the SXR emission, is critical
to properly interpret these broadband measurements.
Our goal with MinXSS observations is to reduce these
SXR spectral differences from factors of 2 or more
down to less than 30%. In addition, MinXSS will
measure solar SXR spectra with higher spectral
resolution of 0.15 keV FWHM, as compared to the
0.6 keV FWHM resolution of the most recent
analogous instrument, MESSENGER SAX.21 The
MinXSS measurements will enable improvements to
solar spectral models, such as CHIANTI and the Flare
Irradiance Spectral Model (FISM).22,23 By using
MinXSS to improve the FISM predictions in the SXR
range, atmospheric studies over the past 30 years will
be possible, such as those for the well-studied
Halloween 2003 storm period, as well as future space
weather events after the MinXSS mission is completed.
Getting this spectral distribution of solar flare energy in
the SXR range is critical as a driver for atmospheric
variations, and will be discussed in another section.
Figure 2: Solar 0.1-7 nm irradiance currently
measured by broadband SXR photometers onboard
NASA’s SORCE and SDO satellites.
The MinXSS data will also help improve understanding
of the physics of solar flares themselves. The 0.5–9 keV
(0.132.4 nm) range observed by MinXSS is rich with
high-temperature spectral lines from coronal plasma
with temperatures from ~5 to 50 million K, which are
greatly enhanced during even small solar flares.
MinXSS will also observe the underlying free-free and
free-bound continua, extending out to 2030 keV,
which can provide an independent diagnostic of the
emitting plasma temperatures. Understanding how solar
flares heat plasma, especially up to many tens of
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million K, is a pressing question in solar physics, and
the MinXSS observations will provide the best spectral
measurements in this energy range to date. e.g., 19,24,25
Observing the variations of spectral lines in comparison
to the continuum will also provide insight into coronal
elemental abundances, particularly for Mg, Si, Fe, S,
and Ar, to help measure abundances and to understand
how they may vary with solar activity and during flares.
Science Objectives Quiescent-Sun Studies
Examples of data analysis and spectral modeling for
two quiescent (non-flaring) solar measurements made
with the X123 aboard the SDO EVE calibration rocket
flights in 2012 and 2013 are provided by Caspi et al.20
One of the tantalizing results from these two 5-minute
observations is that the coronal abundance of certain
elements is different for the quieter SXR spectrum on
June 23, 2012 than the more active (but not flaring) Sun
on October 21, 2013. These abundance differences
suggest that different heating mechanisms occur in the
quiet network versus active regions, and support the
concept that numerous small impulsive events
(“nanoflares”) could be the source of the active region
heating. e.g., 6,26 Identifying the mechanism responsible
for heating the quiet Sun corona to millions of degrees,
while the photosphere below it is only 6000 K, remains
one of the fundamental outstanding problems in solar
physics.27 We anticipate that 13 months of MinXSS
measurements of the solar SXR spectrum will provide
adequate data on active region evolution and several
flares to more fully address these questions on
nanoflare heating. The SXR variability is about a factor
of 1001000 over the solar cycle and can be as much as
a factor of 10,000 for the largest X-class flares;
MinXSS will be able to observe not only small (A- or
B-class flares), but also emission from the truly quiet
Sun, as well.
Science Objectives Improvements to Earth
Atmospheric Models
Energy from SXR radiation is deposited mostly in the
ionospheric E-region, from ~80 to ~150 km, but the
altitude is strongly dependent on the incident solar SXR
spectrum. This wavelength dependence is due to the
steep slope and structure of the photoionization cross
sections of atmospheric constituents in this wavelength
range. The main reason that Earth’s atmospheric cross
section changes so dramatically in this range is due to
the K-edges of O at 0.53 keV (2.3 nm) and of N at 0.4
keV (3.1 nm). Figure 3 shows two different solar SXR
spectra (top) and the result of their absorption in Earth’s
atmosphere (bottom). Although the two solar spectra
are normalized to have the same 0.1–7 nm integrated
irradiance value, their peak energy deposition near the
Earth’s mesopause has a separation of about 5 km. This
separation is considered significant because it is
approximately equal to the scale height at 100 km, it is
critical to E-region electrodynamics, and the mesopause
(the coldest region of the atmosphere) is a critical
transition between the middle and upper atmosphere.
Figure 3: (top) Two examples of CHIANTI model
solar spectra at 0.01 nm resolution, scaled to have
identical 0.1-7 nm integrated energy flux: a Sun with
bright but non-flaring active regions (green), and a
solar flare (red). (bottom) Earth atmospheric
absorption profiles resulting from the two incident
solar spectra in the top plot.
The MinXSS solar SXR spectra are also important to
address outstanding issues concerning E-region
conductance that has an enormous effect on global
electrodynamics and the F-region, especially through
the influence of the equatorial electrojet. One of the
issues concerns the inability of global general
circulation models or detailed process models to
produce enough ionization to agree with the E-region
peak densities from measurements or well-established
empirical models. There appears to be insufficient
energy in the solar spectra used as model input, either
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in the SXR region (especially ~13 nm) or at H Lyman-
β 102.6 nm. The latter has been well quantified by
TIMED and rocket measurements. Thus, the focus on
the solar SXR spectrum may reveal this missing energy
for the E-region. The models could more accurately
describe important phenomena such as the magnitude
and morphology of the equatorial ionization anomalies,
pre-reversal enhancement of the vertical electric field,
and the effects of tidal perturbations on the F-region.
Project History and the Near Future
The MinXSS project began as a graduate student
project in the Aerospace Engineering Sciences (AES)
department at CU and ran through the Spring 2014
semester, with an average of 11 graduate students each
semester. The AES department supported the first year
of the project, which focused on development of the
CubeSat Card Cage design and assembly (see later
section). The second year was supported by the
National Science Foundation (NSF), which enabled the
design and manufacture of the first flight prototype (see
Figure 1 left), which included the structure, command
and data handling (CDH) custom board, electrical
power system (EPS) custom board, custom
motherboard, custom battery pack, and plastic 3D
printed prototypes of the secondary instrument housing
and antenna deployment module. NASA awarded full
funding in the project’s third year to support flight
build, integration, environmental testing, mission
operations, data analysis, and public data distribution.
At the present time, FM-1 has completed environmental
testing and is ready for delivery and launch later in
2015, with deployment expected in early 2016. FM-2
has been built and integrated and is now ready for
environmental testing in late 2015 and launch in mid-
2016. Mason et al. (2015) provides additional details
about the launch vehicles for both FMs and ISS
deployment for FM-1.1
Subsystems
MinXSS has all the standard subsystems for a satellite,
except propulsion. Figure 4 shows a high-level
electrical block diagram for MinXSS. A mechanical
block diagram can be found in Mason et al. (2015).1
The purple boxes in the upper right of Figure 4 show
the Remove Before Flight (RBF) and three rail switches
that disconnect the battery from the system. This
number of switches was required by the NanoRacks
deployer to comply with ISS human safety standards.
Figure 4: MinXSS electrical block diagram.
Motherboard
MinXSS Electrical Block Diagram
8.6 V
Buck
PV1
+Y
8.6 V
Buck
PV2
+X
PV3
-Y
MOSFET 3
MOSFET 2
Battery
COMM
_
+
X123
Detector
External
Power
Supply
MOSFET 1
3.3 V
Buck
5 V
Buck
ADC
INA3221
I
V
I2C
ADC
INA3221
I
V
I2C
ADC
INA3221
I
V
I2C
8.6 V
Buck
E3
E2
CDH
50 mA
Heaters
30 mA
SPS
5 mA
ADCS
1190 - 1940 mW
X123
Electronics
500 mA
Li-1 Radio
700 mA
50 mA
ADC
INA3221
Discharge
I
V
I2C
Fuel
Gauge
MAX17049
E1
I2C
ADC
INA3221
I
V
I2C
V_Batt
ADC
INA3221
Charge
I
V
I2C
MOSFET
MOSFET
MOSFET
ADC
INA3221
I
V
I2C
ADC
INA3221
I
I2C E= Enable
RBF
Rail1 E1
V_Batt
RBF Rail2
E2
V_Batt
RBF
Rail3
Green Tag
E3
V_Batt
TMP100 I2C
TMP100 I2C
EPS Board
I
I2C
I
T
T
T
T
T
ADC
AD 7998
PV Batt
MOSFET
6-8.4V
MOSFET
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A resource breakdown and outline of each subsystem
can be found in Mason et al (2015). To summarize
each: the attitude and determination control system
(ADCS) is a Blue Canyon Technologies (BCT) XACT
unit; the EPS, CDH, battery board, motherboard,
antenna, and SPS & XS instruments were custom
designed and built at CU and LASP; the radio is a
AstroDev Li-1 radio on a custom PCB; the solar cells
are Azur Space 30% efficient, triple-junction GaAs; and
the primary science instrument is an Amptek X123-
SDD.
ADVANCING CUBESAT TECHNOLOGIES AND
LESSONS LEARNED
This section discusses lessons learned that were
incorporated in MinXSS and new lessons from the
development. It also makes calls to the community that
will aid in future CubeSat development efforts.
CubeSat Card Cage
Experience with the PC104 PCB interface on the
Colorado Student Space Weather Experiment (CSSWE)
CubeSat led us away from the card stack design due to
the difficulty in debugging boards once integrated.28,29
Instead, our CubeSat Card Cage design uses a
motherboard/daughterboard architecture that allows any
individual card to be easily removed, and an extender
board optionally inserted to have access to the
daughterboard for probing while still electrically
connected (Figure 5). Additionally, the standard
electrical interface allows boards to be swapped to any
position. MinXSS utilizes a DIN 48-pin connector for
the daughterboard-motherboard interface. This
relatively large connector was chosen for ease of
soldering for new engineering students and because it
easily satisfied the requirements on the number of
necessary pins and mechanical dimensions. In the
future, a higher density connector with potentially more
pins could be chosen to provide a lower mass and lower
volume solution while still providing the flexibility of
the card cage architecture.
3D Printed Parts
MinXSS used 3D printed parts for both prototyping and
flight components. For prototyping, the SPS & XS
housing was 3D printed in plastic twice as the design
iterated, and the solar array hinges were printed in
plastic once. This was done using CU’s Objet 30 printer
with VeroWhitePlus plastic. For flight, these same
components were 3D printed in metal using direct metal
laser sintering at GPI Prototype. The SPS & XS
housing is aluminum with a shot blasted finish. This
finish was very rough and required significant sanding
to get an acceptable surface finish and clean edges. The
solar array hinges are stainless steel with a shot blasted
finish. A minimal amount of sanding was required for
these parts because the requirements were looser and
the finish was slightly better than SPS & XS. The better
finish was likely due to the hinges being a simpler part
that required no filler material during the 3D print
(sintering) process.
As plastic 3D printers become more pervasive,
affordable, and precise, the draw toward using the
resultant parts for flight is becoming stronger. A major
risk that must be addressed is the unknown properties
of these materials, particularly in their response to
vacuum and UV exposure. We would like to see an
open database where specifications based on test results
for common 3D print materials, such as ABS and PLA,
could be accessed.
Figure 5: Prototype CubeSat Card Cage design with
Amptek X123-SDD on top.
Simplification of Solar Panel Fabrication Process
CSSWE used epoxy (Arathane 5753) on the back of
solar cells to adhere them to the solar panel PCBs. This
technique is typical but requires significant assembly
and curing time. MinXSS used double-sided Kapton
tape with acrylic adhesive to adhere solar cells to the
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PCBs. We used a specialized rubber vacuum sealer to
apply pressure to the cells uniformly and meet the
manufacturer’s recommended application pressure.
This reduced the time to produce a solar panel from
three days to one day. To get electrical conductivity
from the back of the solar cell to the PCB, we applied
silver epoxy in large vias behind each cell. We also
tested a new-to-market tape, called Z-Axis tape by 3M,
which is electrically conductive between the adhesive
and back sides and could save the extra step of applying
the silver epoxy or soldering/welding on tabs. We
decided to use Kapton tape for flight because the Z-
Axis tape adhesive was not rated for as wide a
temperature range as the Kapton acrylic adhesive, there
was concern that the Z-Axis tape could not sustain the
high current of the solar cells for as long as solder or
silver epoxy could, and we did not understood the Z-
Axis tape thermal conductivity properties as this
information was not available in the Z-tape
specification sheet.
In the future, we would like to see solar cell
manufacturers adopt a standard form factor compatible
with CubeSats. MinXSS uses 40 mm × 80 mm cells
from Azur Space, which are a great fit within the rail
boundaries of CubeSats (maximum of 83 mm wide and
340.5 mm long for 3U CubeSat). The 80 mm width for
cells provides 1.5 mm margin on each side from the
rails. If the spacing between cells could be reduced to
4.5 mm or less, then there could be 8 Azur Space solar
cells instead of 7 on a 3U panel. Alternatively, if the
height of the cells were changed to be 50 mm instead of
40 mm, then they would be more modular for fitting
one solar cell per 0.5U of the panel length. With six 50
mm x 80 mm cells instead of seven 40 mm x 80 mm
cells, there could be 7% more power per 3U panel.
Pseudo-Peak Power Tracking
We modified the DET EPS design that was inherited
from the CSSWE CubeSat to include an additional
specially selected resistor to create a pseudo-peak
power tracking (PPPT) system. The extra resistor was
chosen to prevent a rapid voltage drop from the solar
cells when the battery attempts to draw a large current,
namely when the battery state of charge is relatively
low right as the spacecraft exits the orbit eclipse.
In the CSSWE and MinXSS EPS design, the output of
the solar panels power 8.6 V regulators that then
provide regulated 8.5 V power directly to the battery
and system. In this DET design, the batteries will
charge up to 8.5 V, and there are no supporting
electronics required to control the battery charging
process. In reality, this simple approach only provides
about 50% of the power intended from the solar panels
when the battery capacity is low. In particular, when the
battery needs more power input (high current) for
charging, the high current draw from the solar cells
results in much lower voltage, following the standard
solar cell current-voltage (I-V) curve. When the solar
panel output voltage goes below the minimum input
voltage level of the 8.6 V regulator, the regulator turns
off. Consequently, the current drops and the solar panel
output voltage increases, and the 8.6 V regulator turns
back on. This results in a high-frequency on-off
regulator oscillation that had the EPS 8.6 V regulators
on for only about 50% of the time during the early part
of the orbit dayside during mission simulations. The
MinXSS solar panels were designed for 80% of peak
efficiency at EOL, but the 50% decrease in power was
an unacceptable power loss for the nominal power
budget.
The solution for MinXSS, without having to redesign or
rebuild the EPS board, was to replace the sense resistor
on the output of the solar panel regulator with a larger
resistance so that the effective current draw out of the
solar panel would be limited and thus would not cause
the regulator to turn off. We refer to this current-
limiting resistor for the solar panels as pseudo-peak
power tracking (PPPT). Figure 6 shows a simplified
version of the PPPT circuit for the MinXSS EPS.
Figure 6: A simplified circuit diagram of PPPT used
for the MinXSS EPS.
The value for this current-limiting resistor was
estimated for the MinXSS power configuration using
Eq. 1. The first term on the right-hand side of Eq. 1 is
the current for the spacecraft load, and the second term
is the current for charging the battery. The spacecraft
load is assumed constant, but the battery charging
current starts off high when the battery voltage is low
and then ramps down to zero when the battery voltage
is the same as the regulator voltage downstream of the
current limiting resistor. The ideal value for the current-
limiting resistor, RCL, is such that it limits the current
out of the regulator, IReg, to be less than the maximum
current, Imax, possible from the regulator (at the peak
power part of the solar panel I-V curve) and when the
battery voltage, Vbatt, is at the lowest allowed level. For
the MinXSS design and configuration, the regulator
voltage, VReg, is 8.5 V, the worst-case system load
(largest power) has 7.0 Ω for RS/C, a battery impedance
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of 0.125 Ω, and a value of 2.8 A for Imax. The goal for
MinXSS was to keep the battery voltage above 7.1 V at
all times, so an RCL of 0.25 Ω is the desired value for
the MinXSS configuration to satisfy Eq. 1. That is, with
this value of RCL, IReg equals Imax when Vbatt equals 7.1
V.
(1)
After the current-limiting resistor was installed into the
EPS, additional mission simulations were run. We
verified that the prediction of the regulator current, IReg,
and the measured battery voltage agreed with the
measured regulator current.
One disadvantage to the PPPT implementation is that
there is additional heating of the EPS board because of
the larger resistance; however, this extra heating peaks
right after exiting eclipse, the precise time when
temperatures are cooler and heating is desired anyway.
For example, the power loss (heating) in the PPPT
current-limiting resistor is estimated to be 2.6 W when
the battery voltage is at its lowest value of 7.1 V,
decreasing to 0.93 W when the battery voltage is at 7.5
V, and reduces to less than 0.1 W once the battery
voltage is above 8.0 V. The primary caveat in the PPPT
design is that resistor tuning must be done a priori, and
is fixed, whereas maximum PPT (MPPT) systems can
tune resistance in real time to maintain the maximum
power point on the solar cell I-V curve. The trade
studies performed for CSSWE and MinXSS resulted in
the selection of a custom DET EPS due to the
simplicity of design. Both teams were unaware of the
consequential loss of power generation at the time of
the original designs. The advantage of the PPPT circuit
is that it is only minimally more complex than DET,
adding little risk for a large benefit.
In the future, we would like to see a standard MPPT IC
for interfacing to common CubeSat battery packs (e.g.,
8.4 V Li-polymer battery packs). We found it difficult
to identify a commercial MPPT IC or proven MPPT
circuit that could be integrated with our system. We
purchased the most promising MPPT IC, a Linear
Technology LT3652, and spent significant time
attempting to integrate it with the MinXSS EPS, but its
intended use prevented proper functioning for our solar
panel and battery configuration.
Importance of Flight-like Testing
Various tests were performed on MinXSS that were
geared toward simulating the orbital environment and
flight-like operations. These included low-external-
torque tests of the ADCS, thermal vacuum with a long-
duration mission simulation, early-orbit end-to-end
communication testing performed several miles away
from the ground station, and detailed battery
characterization of the actual batteries to be flown.
Using a custom-built air-bearing table, we tested the
functionality and performance of the ADCS. This test
simulated an orbital environment with reduced external
torques present. Through this testing, we discovered
that an operational amplifier (op-amp) was preventing
the XACT coarse sun sensor from being properly read
by its internal flight software, and this op-amp was
replaced to resolve this issue. It is unlikely this would
have been discovered otherwise, and may have resulted
in the spacecraft not being able to quickly find or
accurately track the Sun on orbit. Significant effort in
mission operations may have been able to salvage the
mission in that situation, but only minor effort was
required to replace the offending op-amp. Air-bearing
testing requires very careful balancing of the system
and as much reduction of external torques as possible;
e.g., even air flow from building ventilation could limit
the tracking duration while operating on the air-bearing
table. It also requires the computation of moments-of-
inertia specific to the air-bearing-CubeSat system to be
provided to the ADCS for appropriate control to be
implemented. Without such an update to the ADCS
software, the ADCS response is too sluggish (slow) to
confirm that the ADCS is tracking as expected.
Thermal vacuum tests are irreplaceable for determining
if the CubeSat can function in vacuum and for
measuring performance near the operational limits of
components. Through such testing of MinXSS, we
discovered a short in a battery heater that reset the
entire system every few seconds, which only
manifested under vacuum. This was caused by the
battery expansion, which created an unintended
electrical connection between the two nodes of the
heater. Typically, CubeSats are only required to bake
out, not perform a functional thermal vacuum test, but
we highly recommend this test as a process to increase
the success rate of CubeSats.
A 100-hour mission simulation test was performed on
MinXSS during four of the eight hot-cold cycles of the
thermal vacuum testing. A solar array simulator, with
an I-V curve programmed to model the Azur Space
solar cells used on MinXSS, was jumpered into the
MinXSS EPS board. The jumper bypassed the two
deployable solar panels. The output of the solar array
simulator was programmatically cycled in intervals
corresponding to ISS orbit insolation/eclipse periods at
three different β angles. The total orbit period was 93
minutes and the three eclipse periods were 28 minutes
(average β), 38 minutes (β = 0º), and 0 minutes (β >
IReg =VReg Imax RCL
RS/C
+VReg Imax RCL VBatt
RCL +RBatt
Mason 9 29th Annual AIAA/USU
Conference on Small Satellites
76º). We collected power performance data of the entire
system throughout each of these scenarios, and verified
that the PPPT maintained a power positive state through
many orbits. Additionally, this test was used to verify
the functionality of a flight-software commandable flag
to disable power to the X123 during eclipse periods.
This option was introduced into the flight software
early in the project in anticipation of a marginal power
balance. The X123 was chosen for power cycling
because it is the largest consumer of power and because
the primary science target the Sun is not visible in
eclipse. However, this is not the default state in the
mission design as it introduces excessive power cycling
on the primary science instrument; nominal operations
leave the X123 powered on during the entire orbit. As
the spacecraft performance degrades on orbit (e.g., solar
cell efficiency loss), it may become necessary to enable
the X123-eclipse-power-cycling flag. Finally, the 100-
hour mission simulation test included periodic stored-
data downlinking with durations equivalent to the
ground station contacts expected on orbit. The 100-hour
mission simulation test was the most flight-like testing
possible with the facilities available, and greatly
increased confidence in and understanding of the
system as it will behave on orbit. It also ensured that the
flight electronics are likely past the “infant mortality”
phase.
End-to-end testing was also performed on MinXSS to
verify functionality of the full communication pipeline.
The spacecraft was taken several miles away to a
position in the line-of-sight of the ground station, and
early-orbit commissioning tests performed. This
boosted confidence in several areas: that we would
meet the NanoRacks requirement of not deploying the
MinXSS antenna or solar arrays in the first 30 minutes
after deployment from the ISS, that those deployments
would be successful, that communications could be
established after antenna deployment, and that our
ground software commissioning scripts could
autonomously perform telemetry verification and
commanding.
Significant battery testing was performed to comply
with requirements flowed down from NASA Johnson
Space Center through NanoRacks to all CubeSats going
to the ISS. These requirements are in place to protect
astronauts on the ISS and far exceed the standard
CubeSat requirements in the Cal Poly CubeSat Design
Specification. Nevertheless, we recommend that all
CubeSats perform several of these tests, if only to better
understand the actual batteries to be flown, i.e., not just
batteries from the same lot or of the same type. We
found the following to be the most useful tests: visual
inspection for dents or leaks, measuring the open circuit
voltage of the fully configured battery pack, recording
voltage, current, and temperature through three
charge/discharge cycles, measuring the voltages at
which overcharge and over-discharge protection
activated and deactivated, and measuring mass before
and after undergoing vacuum. Given availability of the
equipment to perform these tests and measurements, it
took approximately two weeks to complete this testing
for each battery pack. Much of that time was dedicated
to setup, waiting for charge cycles to complete, and
interpretation of the results. Additional tests were
required for astronaut safety on the ISS, but we would
consider them to be extraneous for non-ISS CubeSat
missions. These include measuring of the physical
dimensions of each battery, measuring the closed circuit
voltage of the fully configured battery pack, measuring
the time to trigger short-circuit protection and
maintaining the short for three hours to verify the
protection remains enabled, and doing a dedicated
vibration test at five frequencies and strengths up to
9.65 grms on all three axes, with voltage measurements
between each axis. These additional tests took several
weeks of additional time and planning, particularly in
the design, manufacturing, and modification of
components to support vibration testing.
Importance of a Second CubeSat Unit
The fabrication of two identical sets of hardware in
parallel is much less expensive than the same
development in series, particularly if the start of the
development for the second set is delayed by months or
years. Small projects tend to have less stringent
requirements on documentation, so details can be
forgotten and lost in the time between two sets of flight
hardware developed in series. Having two sets of
hardware enables the development and testing of flight
software while other activities proceed in parallel.
Importantly, parallel development also enables the
replacement of a subsystem if a problem is found,
which is critical when schedules are tight. This was the
case for MinXSS when the battery heater short was
discovered in FM-1 at the initial pump-down for its
thermal vacuum test. We were delayed half a day to
swap the battery pack out with FM-2, which did not
have the same issue, as compared to the weeks of delay
that would have been introduced if an entirely new
battery pack had to be assembled and tested. Finally,
having a second flight unit allows for debugging of
hardware and software after delivery and launch of the
first flight unit.
Low-cost Mitigation of Radiation Issues for
Electronics
The CubeSats developed at CU and LASP have
generally used industrial-grade (automobile) electronic
parts because those parts have wider operating
Mason 10 29th Annual AIAA/USU
Conference on Small Satellites
temperature ranges. Typically, the automobile-grade
ICs cost $10 as compared to $2 for standard
commercial ICs, but this additional cost is outweighed
by the significant benefits of the higher-grade
components. For example, the number of uncorrupted
SD card write cycles can be improved by a factor of
10100, and the operational temperature range
expanded by purchasing a $70 4 GB hardened SD-card
instead of a $4 standard SD-card. The total cost impact
on MinXSS for these industrial-grade electronics parts
is only a few thousand dollars, a small fraction of the
total budget, but it significantly improves the potential
for a longer mission life. While our intention was to
have electronics that could operate over a wider
temperature range, automobile-grade parts may also
help with radiation tolerance of the electronics. Two
MinXSS prototype CDH boards were radiation tested,
one to 10 kRad and another to 25 kRad; both boards
survived. It is not clear if industrial-grade parts made a
difference or not for passing the harder radiation test;
nonetheless, it is only a small cost increment to use the
higher-grade parts.
SUMMARY AND CONCLUSIONS
MinXSS is a 3U CubeSat developed at CU and LASP
with two flight models: one to be launched in late 2015
and the second to launch in mid-2016. This
development leveraged heritage from the highly
successful CubeSat, CSSWE, which was also
developed at CU and LASP.30,31,32 The primary science
objective of MinXSS is to fill a critical spectral gap in
solar measurements currently made by large satellite
missions at 1/100th their typical cost. All standard
satellite subsystems are present in MinXSS, except
propulsion, packaged in a volume that can fit in a
breadbox. Many of these subsystems were custom
developed by CU and LASP (e.g., CDH, EPS, SPS &
XS, structure), primarily by graduate students with
professional mentorship, and other subsystems were
purchased from commercial vendors (e.g., flight radio,
ADCS, primary science instrument).
In the future, 6U, 12U, and 27U CubeSat standards will
open up even more science capabilities by allowing for
larger and more sophisticated instruments. Standardized
buses that can be commercially procured are now
becoming available. Typically 12U in size, this leaves
ample mass and volume to be used for the science
payload. As X-band transmitters and LEO-accessible
global network communications, such as GlobalStar,
become available in the near-term, it will also be
possible to expand data downlink capabilities. This
increase in data volume is a critical need for science
that involves imaging, as even a single image from a
small camera would take hours to downlink at the
CSSWE / MinXSS rate of 9600 bps. We note that
active pixel CMOS array detectors provide one
alternative mitigation strategy for this, if the entire
image does not need to be downlinked. Finally,
CubeSats will enable science that was not conceivable
with large, monolithic spacecraft. For the same cost as a
NASA Small Explorer, a constellation of dozens of
CubeSats could be put into orbit to obtain simultaneous
measurements over a wide spatial distribution. These
novel data will enable new scientific observational
analyses and provide new constraints to physical and
empirical models.
ACKNOWLEDGEMENTS
This work was supported by NASA grant
NNX14AN84G and NSF grant AGSW0940277, as well
as the University of Colorado at Boulder Aerospace
Engineering Sciences Department. We would like to
specially thank the many professional scientists and
engineers that provided feedback at reviews and
mentorship to students, and to thank the many students
who made contributions above and beyond what was
required.
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