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On-orbit verification of space solar cells on the CubeSat MOVE-II

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Several promising multi-junction solar cell concepts for space applications are currently under development worldwide. On-Orbit Verification on CubeSats is a cost-efficient method to gain data on critical hardware early in the design validation process. The MOVE-II CubeSat will be used for the verification of novel 4-6 junction solar cells. With a footprint of 10x10 cm², the payload consists of one full size solar cell (8x4 cm²) and up to 7 positions (each 2x2 cm²) for corresponding isotype solar cells. The measurement electronics is based on commercial off-the-shelf hardware. MOVE-II is planned to launch in early 2018 into a 500-550 km sun-synchronous orbit.
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On-Orbit Verification of Space Solar Cells on the CubeSat MOVE-II
Martin Rutzinger1,2, Lucas Krempel1, Manuel Salzberger1,2, Mario Buchner2, Alexander Höhn1,
Maximilian Kellner1, Katja Janzer1, Claus G. Zimmermann2, Martin Langer1
1 Technical University of Munich, Germany
2 Airbus DS GmbH, Munich, Germany
Abstract Several promising multi-junction solar cell
concepts for space applications are currently under development
worldwide. On-Orbit Verification on CubeSats is a cost-efficient
method to gain data on critical hardware early in the design
validation process. The MOVE-II CubeSat will be used for the
verification of novel 4-6 junction solar cells. With a footprint of
10x10 cm², the payload consists of one full size solar cell (8x4 cm²)
and up to 7 positions (each 2x2 cm²) for corresponding isotype
solar cells. The measurement electronics is based on commercial
off-the-shelf hardware. MOVE-II is planned to launch in early
2018 into a 500-550 km sun-synchronous orbit.
Index Terms — CubeSat, on orbit verification, multi-junction
solar cell, degradation.
I. INTRODUCTION
The majority of today's satellites is powered by photovoltaic
generators. New multi-junction solar cells (MJSC) for space
application are currently under development. Amongst the
most promising concepts are 4-6 junction MJSC’s [1]-[3]. In
addition to terrestrial and laboratory-based characterization
methods [4],[5], this highly dynamic technology development
requires accurate, flexible and low cost on-orbit verification
(OOV) methods. This paper introduces the design of such an
OOV mission using the CubeSat MOVE-II.
Since their introduction in 1999, CubeSats [6] have evolved
from purely educational missions to spacecraft with a broad
variety of scientific and commercial applications. OOV of
critical hardware has been the objective of multiple CubeSat
missions in the past [7][10]. Thereby, the good cost-
efficiency of CubeSats allows the OOV at an early stage in
product development and at a relatively low Technology
Readiness Level (TRL) [11]. The Munich Orbital Verification
Experiment (MOVE) satellite program of the Technical
University of Munich was initiated in 2006 with the ambition
of building a single-unit CubeSat verification platform, called
First-MOVE. The main goal of the program since then has
been the hands-on education of undergraduate and graduate
students. First-MOVE was launched in late 2013 and operated
in space for a month [12]. Currently, the second satellite of the
program, called MOVE-II, is under development [13].
MOVE-II is planned to launch in early 2018 into a 500-
550 km sun-synchronous orbit (SSO).
Besides the educational goals of the MOVE-II program, the
technical mission objective is to measure the current vs.
voltage (I-V) characteristics of novel multi-junction solar cells
(MJSC) in orbit. The main reason for the OOV of novel
MJSC's is to gain flight experience for these cells. MOVE-II
will provide complimentary data to earth bound [4],[5] and
high altitude balloon or aircraft based [14] experiments. The
main advantage of OOV is long time exposure to space
environment, especially the solar air mass zero (AM-0)
spectrum. This spectrum cannot be exactly reproduced with
earth bound sun simulators and even experimental data
obtained from high altitude experiments has to be corrected
for residual atmospheric absorption [14].
The MOVE-II platform is designed for maximum
flexibility, so that the exact type of MJSC, either with 4, 5, or
6 junctions, will only have to be fixed in late 2016, based on
the technological availability at that time. In addition to one
MJSC, also 4-6 corresponding component (isotype) solar cells
which have the same optical properties as the MJSC but only
one activated p-n junction will be characterized. This provides
additional information about each individual subcell.
As depicted in Fig. 1, the top plate (red area) of the
CubeSat (10x10 cm²) will be utilized for those payload cells.
Fig. 1 Rendering of the MOVE-II CubeSat: the payload area for
novel MJSC's is highlighted by the red frame on the Zenith deck.
II. MISSION AND PAYLOAD DESIGN
A. CubeSat Platform MOVE-II
MOVE-II will be designed as a 1 Unit CubeSat
(10x10x10 cm³), with a total mass of 1.3 kg. The satellite will
incorporate a UHF/VHF transceiver for telemetry and low-to-
mid data rate transmission. Furthermore, an experimental S-
Band transceiver will enhance the capabilities of the satellite
towards higher data rates. The achievable data rate on the S-
Band link is expected to be more than 1 MBit/s. It is planned
to use the downlink capability in both bands for payload data
transmission.
The attitude determination & control system (ADCS) is
based on magnetic actuation and will allow a pointing of the
satellite with accuracy better than ±10°. In its nominal mode,
the satellite will operate in a position where the payload is
pointing in zenith direction (Fig. 1).
B. Payload Architecture
The payload assembly on the Zenith deck consists of one
full size MJSC (8x4 cm²) and 5 to 7 open positions (each
2x2 cm²) for corresponding isotype (component) solar cells
and additional experimental cells (Fig. 2). The solar cells are
mounted on the outboard side of a printed circuit board (PCB).
The inboard side of the PCB contains all necessary
measurement electronics.
C. Measurement Electronics
Each solar cell is contacted with a 4-point connection
(Kelvin-connection) with separate contacts for the current path
and the voltage measurement. The solar cell voltage UC is
measured directly between positive and negative contact of the
solar cell. The current is measured as voltage drop US over a
precision shunt resistor    with low temperature
coefficient.
The current vs. voltage (I-V) curve sweep between ISC and
VOC is performed by varying the gate voltage of a MOSFET
which serves as a variable electronic load. The voltage sweep
is controlled through a Digital Analog Converter (DAC). Both
voltage signals, cell voltage and shunt voltage, are digitalized
with 24-bit Analog Digital Converters (ADC’s). The selected
COTS parts and their tested radiation hardness are
summarized in Table 1.
The circuit diagram for one payload solar cell is shown in
Fig. 3.The temperature of each solar cell is measured with a
digital sensor which is in direct thermal contact with the back
surface of the solar cell.
The I-V curve measurement accuracy is determined mainly
by the temperature stability of the current sensing shunt
resistor and by the temperature- and radiation-stability of
the ADC.
A prototype with all selected components has been built and
tested with 4J-UMM solar cells and corresponding component
cells J1-J4 from AZUR SPACE Solar Power GmbH [3]. Fig. 4
shows a comparison of the I-V curves of a 4x8 cm² 4J-UMM
solar cell and 2x2 cm² component (isotype) cells measured
with a calibrated sourcemeasurement unit (SMU) and with
the MOVE-II prototype. All I-V curves were measured under
an uncalibrated AM-0 spectrum at room temperature. The
results obtained from this comparison show that the prototype
can yield accuracy better than 0.5% in voltage and current
measurement. A sufficient amount of data points is collected
to allow exact fitting of diode parameters.
Fig. 2 Layout and solar cell positions on the MOVE-II payload in
the configuration with one 4x8 cm² (L1) and 2x2 cm² (S1 S7) solar
cells.
Fig. 3 Circuit diagram for one payload cell of MOVE-II.
TABLE 1:
SELECTED COTS PARTS FOR THE MEASUREMENT
ELECTRONICS.
Component
Part number
Rad. Hardness
DAC
TI DAC-7512E [15]
30 krad [16]
ADC
AD 7714 [17]
10 krad[18]
MOSFET
Fairchild FDS 8870 [20]
Not tested
Shunt resistor
Vishay CSM2512 [21]
Not tested
Temp. sensor
ADT7310 [22]
Not tested
Complimentary data, such as sun angle and earth visibility,
are collected by the ADCS subsystem and also included in the
transmitted data. This input will be used for on-ground post-
processing and for the investigation of the effects of sun
incidence angle and earth albedo on the solar cell
performance. The seasonal sun-earth distance has to be
corrected by Eq. 1 [23]
      (1)
where r is the sun-earth distance in astronomical units (AU).
If I-V curves can be obtained at different temperatures it
will be possible to extract the temperature coefficients
according to Eq. 2.
 
    (2)
The quantity X represents   
, or
 .
III. MISSION SIMULATION
Preliminary simulation results show the feasibility of the
OOV-experiment on MOVE-II.
A. Lifetime Analysis
The maximum lifetime of MOVE-II was calculated using
the STELA tool, version 2.6.1, from CNES [24]. The 500 km
SSO orbit resulted in an average lifetime of 3.6 years, versus
7.6 years for the 550 km orbit height (worst case = minimum
drag area of 10x10 cm²). Therefore, both the feasibility of an
OOV of solar cells due to degradation effects over time and
compliance to the space debris mitigation guidelines [25] are
ensured.
B. Radiation Degradation Analysis
We calculated the expected electron and proton radiation
damage on triple junction solar cells using the software
SPENVIS 4.6.7 [26] for a circular 550 km SSO with 97.5°
inclination and 7.6 years lifetime. With a 100 µm thick cover
glass, a 1 MeV End-of-Life (EOL) equivalent electron fluence
      can be calculated. For well-
studied triple junction solar cells, this fluence typically leads
to current, voltage and power remaining factors > 99% [28].
Therefore, no useful degradation analysis can be expected in
the given orbit for solar cells with cover glass.
For a purely experimental study of degradation effects, one
additional 2x2 cm² MJSC will be mounted as bare cell, i.e.
without a cover glass. In this case the calculated equivalent
fluence  is increased to    cm-2. This
high dose mainly arises due to the lack of proton shielding,
especially the low energy part of the proton spectrum which
normally is completely absorbed in the cover glass. In triple
junction cells, such a dose would result in significant
degradation with a power-remaining factor of 
[28]. In this orbit, though, the particle spectrum is dominated
by low energy protons and most of them would be absorbed in
the top junction. Therefore the bare MJSC degradation
behavior is expected to be dominated by the degradation of the
top subcell. Degradation due to low energetic protons has been
studied with mono-energetic accelerator protons under normal
incidence [29] and for omnidirectional incidence [30].
However, no on-orbit verification of these ground based
experiments is available up to now.
C. Sun Pointing Mode
A continuous sun-pointing mode of a CubeSat can be
achieved with a magnetically actuated spacecraft [31]. Thus,
although MOVE-II will operate nominally nadir-pointing, it is
planned to also maneuver in sun-pointing mode as a stretch
goal for the mission. The relevance of the technical data
collected in sun-pointing mode and thus, peak power can be
improved under the expected mission operation scenarios.
Fig. 4 (a) and (b) Comparison of I-V curves of a 4J-UMM solar cell
and corresponding component cells J1-J4 from AZUR SPACE Solar
Power GmbH [3]. The I-V curves were measured with a source
measurement unit (solid lines) and the MOVE-II prototype (open
squares) under an uncalibrated AM-0 sun simulator.
D. Thermal Simulation
To ensure the compliance with the operational limits of the
solar cells and the measurement electronics, thermal
simulation and experiments have been carried out. Thermal
modeling has to deal with various uncertain parameters, such
as unknown conductance through screws or adhesive contacts.
Thus, to achieve reliable thermal simulations, correlation with
actual experiments is important.
For the MJSC experiment, a thermal prototype of an
aluminum CubeSat frame, equipped with two inner dummy
PCBs, and an experimental PCB with two 30.2 cm² cells, was
built. This setup was equipped with ten Pt-100 thermocouples,
observing temperatures directly on the solar cells, the
electronics side, the aluminum structure, and an inner dummy
board.
A thermal vacuum experiment was performed, insolating
the prototype with 1.05 × AM-0 solar intensity for 38 minutes,
and tracking the temperature development until an
approximate steady state was achieved. The chamber walls
were not cooled during testing, and heated up to +42 °C
background temperature. Due to lack of convection, panel
temperatures rose to +147 °C during insolation. Heat
conduction through the Zenith board pushed electronic
temperatures up to +130 °C, while inner PCBs of the satellite
kept more moderate temperatures of max. +77 °C. Although
these values significantly exceed the maximum operation
temperature for solar cell electronics on the Zenith board
(+105 °C), the experimental setup represents an unrealistic
absolute hot case. In space, background temperatures of
 °C will prevent such overheating scenarios, due to more
effective radiative heat exchange.
The thermal vacuum experiment provided important data
for correlating a numerical thermal simulation, performed with
ESATAN [32]. The numerical model represents all important
structural features and screw joints of the experimental
prototype. Contact conductance and aluminum surface
properties were varied, until modelled temperatures
sufficiently met experiment observations. The correlated
model predicted interior temperatures with ± 5 °C accuracy,
and solar cell temperatures with larger negative deviation.
The correlated model was used to predict temperatures on a
550 km, 11 h Local Time of Descending Node (LTDN) SSO
orbit. As the test prototype lacked side panels and wings for
additional solar cells, these elements have been added to
create a realistic MOVE-II model in space. Representative
values for contact conductance from the prototype model have
been transferred to all additional screws and material contacts.
Fig. 5 depicts the temperature distribution of the MOVE-II
model during orbital flight. In a Zenith pointing configuration,
temperatures are much lower than in the thermal vacuum
experiment, due to lower background temperatures and shorter
orthogonal insolation of the panels. Peak temperatures of
MJSC’s experiment electronic reach only +63 °C, well within
the designed temperature window. Lowest temperatures on the
electronics board during eclipse reach 3 °C. Thus, the
preliminary results of the thermal simulation conclude safe
operational conditions for the solar cell electronics.
IV. SUMMARY
The CubeSat MOVE-II will carry a novel 4-6 junction solar
cell and corresponding component cells intended for future
space application with the purpose of measuring their
performance and degradation behavior in low earth orbit. The
developed measurement electronics was shown to be capable
of measuring current vs. voltage curves with high precision.
The launch of MOVE-II is planned for early 2018.
ACKNOWLEDGEMENT
The authors would like to thank AZUR SPACE for their
support and close collaboration.
The authors acknowledge the funding of MOVE-II by the
Federal Ministry of Economics and Energy, following a
decision of the German Bundestag, via the German Aerospace
Center (DLR) with funding grant number 50 RM 1509.
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... Furthermore, MOVE-II should demonstrate a magnetorquer based 3-axis stabilized bus[4]for future science and technology verification missions and should be capable of processing and downloading large amounts of data using the in-house developed Nanolink protocol[5]. As the scientific payload, the degradation of a 4junction solar-cell prototype will be measured in orbit[6]. The MOVE-II satellite is depicted inFig. ...
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Changes due to design flaws impose major costs, delays and high risks on any spaceflight project. The later the change, the riskier and more expensive it is. System changes due to failures detected during spacecraft assembly are usually one of the last hardware flaws to be found and therefore impose major risks on the overall project. Traditionally, this is overcome by metal or wooden mock-ups early in the process. However, to respond to design changes in a fast manner and to properly explore the remaining options by building multiple full size mock-ups in a short time interval, rapid prototyping was used by the authors. This paper provides lessons learned of the Munich Orbital Verification Experiment II (MOVE-II), related to rapid prototyping technologies used during the development phase. MOVE-II is the second CubeSat mission of the Chair of Astronautics at the Technical University of Munich (TUM). Early in the design process, a 3D printed structural model of the CubeSat was built to verify the CAD model, the assembly strategy, and to track down potential system level design deficiencies. By doing so, minor and major flaws concerning integration of the satellite were found in an early project phase. Furthermore, multiple design alternatives were 3D printed during the development process, not only exploring different solutions but also defining cable paths and cable lengths and evaluating the corresponding assembly process. In difference to traditional methods, 3D printing allows for a shorter implementation time span of different design options. In addition it was possible to conduct dress-rehearsals of the integration procedure early on in order to save time in later project phases, and without potentially harming expensive hardware. Due to the early integration of the prototype, Ground Support Equipment (GSE) and specific tools could be defined ahead of time. The biggest non-technical benefit was, that the physical model simplified communication of problems and possible configurations as well as introductions to the system. Display material was always available for the developing team, either for presentations of the project or for recruitment of new team members. The paper concludes with a brief assessment of the limitations of rapid prototyping technologies for risk reduction and process acceleration. Assessment of mechanical functionality as well as mechanical fits are limited due to production tolerances. Therefore the deployment mechanism of MOVE-II could not be tested sufficiently. Finally, future improvements are shown for upcoming CubeSat missions of the TUM.
... As payload of MOVE-II, the performance and degradation of novel solar cells in outer space shall be investigated. 6 The launch into a 575 km sun-synchronous orbit (SSO) is scheduled for early 2018. The assembled engineering model (EM) of MOVE-II is illustrated in Figure 1 ...
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MOVE-II (Munich Orbital Verification Experiment) will be the first CubeSat of the Technical University of Munich (TUM) utilizing a magnetorquer-based active attitude determination and control system (ADCS). The ADCS consists of six circuit boards (five satellite side panels and one central circuit board in satellite stack), each equipped with a microcontroller, sensors and an integrated coil. The design enables redundancy and therefore forms a fault-tolerant system with respect to sensors and actuators. The paper describes the hardware implementation, algorithms, software architecture, and first test results of the integrated ADCS on the engineering unit. A possibility to upgrade and extend our software after launch will enable further research on new and innovative attitude determination and control strategies and distributed computation on satellites. The MOVE-II flight unit is in the integration and test phase with an intended launch date in early 2018.
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MOVE-II is a student-built satellite due for launch early 2018. Its attitude determination & control system (ADCS) uses magnetorquers and consists of one central Mainpanel processing the control algorithms and five peripheral Sidepanels containing the sensors and actuators. In this thesis, a hardware-in-the-loop (HiL) environment is built and operated to verify the functionality of the Mainpanel in a flight configuration. The HiL-simulation contains models of the space environment, worst-case disturbances, spacecraft dynamics, the magnetorquers, and the sensors including worst-case noise levels. The simulation outputs modeled sensor data to the Mainpanel and reads back the magnetorquer commands of the Mainpanel. This way, the test runs are not affected by the disturbances that are present with physical ADCS testing rigs like Helmholtz cages and air bearings. The ADCS control loop is analyzed and a device for interfacing the Mainpanel to the simulation is designed, manufactured, and programmed. The testing environment is automated, so it can change the simulation parameters programmatically. Both the detumbling and the sunpointing controller have been successfully verified in a worst-case simulated space environment. 32 test runs with different controller parameters, different sensor characteristics, and different environmental parameters cover a wide range of conditions that might be encountered during the MOVE-II mission. A sensitivity analysis shows the reaction of the sunpointing controller to modelling inaccuracies. The simulation estimates the power budget with greater accuracy than previous estimation techniques employed for the MOVE-II mission. During sunpointing, the satellite is power-positive in nominal mode. But the power budget in safe mode where the satellite is tumbling could not be verified to be positive. Therefore, the capability of the satellite and especially its ADCS to recover from low battery situations is analyzed. Further tests to increase confidence in the ADCS and work on a fitting environment to analyze telemetry retrieved from the satellite are suggested to assure correct operation of the ADCS during flight.
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