ArticlePDF Available

Design, Implementation, and Operation of a Small Satellite Mission to Explore the Space Weather Effects in LEO


Abstract and Figures

Ten-Koh is a 23.5 kg, low-cost satellite developed to conduct space environment effects research in low-Earth orbit (LEO). Ten-Koh was developed primarily by students of the Kyushu Institute of Technology (Kyutech) and launched on 29 October 2018 on-board HII-A rocket F40, as a piggyback payload of JAXA’s Greenhouse gas Observing Satellite (GOSAT-2). The satellite carries a double Langmuir probe, CMOS-based particle detectors and a Liulin spectrometer as main payloads. This paper reviews the design of the mission, specifies the exact hardware used, and outlines the implementation and operation phases of the project. This work is intended as a reference that other aspiring satellite developers may use to increase their chances of success. Such a reference is expected to be particularly useful to other university teams, which will likely face the same challenges as the Ten-Koh team at Kyutech. Various on-orbit failures of the satellite are also discussed here in order to help avoid them in future small spacecraft. Applicability of small satellites to conduct space-weather research is also illustrated on the Ten-Koh example, which carried out simultaneous measurements with JAXA’s ARASE satellite.
Content may be subject to copyright.
Design, Implementation, and Operation of a Small
Satellite Mission to Explore the Space Weather Effects
in Leo
Isai Fajardo 1,*,† , Aleksander A. Lidtke 1,† , Sidi Ahmed Bendoukha 2,
Jesus Gonzalez-Llorente 1, Rafael Rodríguez 1, Rigoberto Morales 1, Dmytro Faizullin 1,
Misuzu Matsuoka 1, Naoya Urakami 1, Ryo Kawauchi 1, Masayuki Miyazaki 1,
Naofumi Yamagata 1, Ken Hatanaka 1, Farhan Abdullah 1, Juan J. Rojas 1,
Mohamed Elhady Keshk 1, Kiruki Cosmas 1, Tuguldur Ulambayar 1, Premkumar Saganti 3,
Doug Holland 4, Tsvetan Dachev 5, Sean Tuttle 6, Roger Dudziak 7and Kei-ichi Okuyama 1
1Department of Applied Science for Integrated Systems Engineering, Kyushu Institute of Technology,
1-1 Sensui, Tobata, Kitakyushu, Fukuoka 804-8550, Japan; (A.A.L.); (J.G.-L.); (R.R.); (R.M.); (D.F.); (M.M.); (N.U.); (R.K.); (M.M.);
(N.Y.); (K.H.); (F.A.); (J.J.R.); (M.E.K.); (K.C.); (T.U.); (K.-i.O.)
2Satellite Development Center CDS, POS 50 ILOT T 12 BirEl Djir, Algerian Space Agency, Oran 31130,
3Prairie View A&M University, Prairie View, TX 77446, USA;
4Holland-Space LLC, Houston, TX 77059, USA;
5Space Research and Technology Institute-Bulgarian Academy of Sciences, Sofia 1113, Bulgaria;
6Sigma Space Systems, Canberra 2905, Australia;
7University of New South Wales, Canberra campus, School of Engineering and Information Technology,
Canberra 2600, Australia;
These authors contributed equally to this work.
Received: 14 August 2019; Accepted: 19 September 2019; Published: 27 September 2019
Ten-Koh is a 23.5 kg, low-cost satellite developed to conduct space environment effects
research in low-Earth orbit (LEO). Ten-Koh was developed primarily by students of the Kyushu
Institute of Technology (Kyutech) and launched on 29 October 2018 on-board HII-A rocket F40, as a
piggyback payload of JAXA’s Greenhouse gas Observing Satellite (GOSAT-2). The satellite carries a
double Langmuir probe, CMOS-based particle detectors and a Liulin spectrometer as main payloads.
This paper reviews the design of the mission, specifies the exact hardware used, and outlines the
implementation and operation phases of the project. This work is intended as a reference that other
aspiring satellite developers may use to increase their chances of success. Such a reference is expected
to be particularly useful to other university teams, which will likely face the same challenges as the
Ten-Koh team at Kyutech. Various on-orbit failures of the satellite are also discussed here in order to
help avoid them in future small spacecraft. Applicability of small satellites to conduct space-weather
research is also illustrated on the Ten-Koh example, which carried out simultaneous measurements
with JAXA’s ARASE satellite.
Aerospace 2019,6, 108; doi:10.3390/aerospace6100108
Aerospace 2019,6, 108 2 of 38
space environment; plasma; radiation; single events; particle detectors; Langmuir probe;
magnetometer; rapid spacecraft development; small-satellite; Ten-Koh
1. Introduction
Space radiation affects satellites by introducing anomalies such as single event effects (SEE),
component degradation due to ionizing radiation dose, and surface and internal charging.
Understanding the radiation environment is, therefore, important in order to design satellites that
can withstand the possible anomalies. Space radiation sources include galactic cosmic rays (GCR),
solar energetic particle (SEP) events, high energy particles trapped in the Earth’s magnetic field and the
continuous radiation background. In addition, the low-Earth orbit (LEO) region where most satellites
reside is subject to unknown mechanisms that provide it with unpredictable energy variability in the
spectra of particles [
]. This variability is particularly poorly understood for electrons, which can
appear with energies higher than expected.
The consequences of the presence of high-energy electrons, protons and ions for spacecraft
developers vary depending on each mission design, epoch and class. The spacecraft design should
account for the effects of ionizing radiation, as well as charging and discharging effects on satellite
surfaces. For manned missions, mission duration [
] and the associated life support systems need to
be tuned to account for this unpredictability of particle populations.
In recent years, different missions have been launched to explore the near-Earth region in order
to provide direct measurements of charged particles, plasma, and the magnetosphere. Missions
such as the Van Allen Probes (RBSP) [
], MMS [
satellite [
], Proba-2 [
] and Swarm [
] from ESA have been launched into specific orbits to
study space radiation around the Earth [
]. In addition to the mentioned missions with launch
masses of hundreds of kilograms, dedicated small satellites offer an important complement to the
measurements because they enable a wider, more comprehensive view of the space environment thanks
to their reduced development time and cost. This class of satellites leverages commercial off-the-shelf
(COTS) components to bring cost and time savings at the expense of an increased risk of failure.
The small-satellite Ten-Koh has been developed to demonstrate the feasibility of providing space
environment measurements with such low-cost platforms, as well as to provide readily usable data.
This paper provides a comprehensive overview of the Ten-Koh mission design, presents its
preliminary results and discusses the applicability of small, low-cost satellites to conduct space
weather research. The specific objectives of the mission are described next, followed by a summary
of the satellite platform and payloads. The main challenges in the development of the satellite are
then summarized and the adopted solutions to them described. Even though the mission has been
successful, which is demonstrated by briefly presenting data obtained in-orbit, it has suffered a number
of failures at component, assembly and subsystem level. These are described in detail to help improve
the design and implementation of future small satellites that take advantage of the COTS components.
2. Materials and Methods: Ten-Koh Mission Description
2.1. Ten-Koh Project Objectives
The Ten-Koh satellite mission described in this paper has the following primary objectives:
1. To characterize the plasma environment around a spinning spacecraft.
To detect MeV-range electrons in LEO and investigate the space environment in the presence of
an extreme low solar activity.
To investigate the change of physical properties of LATS (Lightweight Ablator series for
Transfer vehicle) and CFRP (Carbon Fiber Reinforced Polymer) material samples exposed to the
space environment.
Aerospace 2019,6, 108 3 of 38
Involving students in satellite development, manufacturing, testing and operations is an important
part of their curriculum and enhances their education. Providing this involvement has been set as a
secondary mission objective of Ten-Koh.
Ten-Koh also provided a flight opportunity for two technology demonstration payloads,
which constituted another secondary objective of the mission. These payloads are the Thermal
Switch developed at Sigma Space Systems, which flight-proves a novel design of a switchable passive
thermal switch [
], and the Ultracapacitor Experiment that aims to quantify the performance of an
ultracapacitor as a satellite energy storage device. These two experiments are not discussed in detail in
this paper because they are not directly related to particle physics and they do not address the primary
mission objectives. However, their location within the Ten-Koh architecture is presented for the sake
of completeness.
From the mentioned primary objectives, the novelty of the Ten-Koh satellite derives from
observations of space radiation and ionosphere environment at the end of the 24th Solar cycle, with
an increase in GCR flux in higher latitudes. This is associated with Primary Objectives 1 and 2.
As for Primary Objective 3, resin composites are expected to deteriorate when exposed to the in-orbit
conditions, with the deterioration rate dependent upon the current state of the LEO environment.
Satellite observations of this environment combined with measurements of the degradation of
composite materials make the Ten-Koh spacecraft unique, especially in the class of small satellites.
This paper focuses on the space radiation and its direct effects on electronics, not material
degradation. Thus, the focus is placed on Primary Objectives 1 and 2 while others, notably Primary
Objective 3, will be discussed in dedicated publications due to their specific nature. However, the
accommodation of the instruments required by Primary Objective 3 is described here for the sake of
2.2. Satellite Platform
Ten-Koh was developed over a period of 16 months mainly at Kyushu Institute of Technology
(Kyutech), while some payload instruments were designed and developed in parallel in Australia,
Bulgaria, and the USA. Ten-Koh is based on the platform of a previous small deep-space probe
Shinen-2 [
], developed and launched by Kyutech in 2014. The main structure of Ten-Koh is composed
of a CFRP composite shell with a rigid internal load-bearing structure made of aluminum alloy
(Al 6061-T6).
The satellite platform was based on several components and architecture topologies used by
Shinen-2. However, the design and integration of most of the platform subsystems has been completely
redefined in order to suit the new mission objectives. The main goal when adapting the Shinen-2
platform architecture for the purpose of Ten-Koh was to reuse as many of the heritage electronic
components as possible in order to reduce the risk of component failure and thus increase system
reliability. The development time was also shortened by reusing the same structure, which assured
both successful environmental testing and removed the need for devoting a significant effort to design
a new structure. Figure 1shows the Ten-Koh satellite in its in-orbit configuration.
The satellite platform is formed of the following subsystems:
OBC—on-board computer and data handling subsystem;
COMM—communication subsystem;
EPS— electrical power subsystem; and
ADS— attitude determination subsystem.
Aerospace 2019,6, 108 4 of 38
Figure 1.
Ten-Koh satellite flight model configuration once in orbit. The black external structure is
made of CFRP. (
) Computer-aided design rendering on the left, flight model photograph on the right.
(b) The envelope allowed by the launch vehicle was 500 mm ×500 mm ×500 mm.
Due to the short time allowed for the development and limited resources, attitude control was not
included in the satellite and the use of any mechanisms was discarded, even though they were being
considered until the preliminary design review (PDR). A passive attitude control based on permanent
magnets was not possible due to the influence on the CPD and magnetometer readings. Figure 2
shows the block diagram of the satellite platform subsystems and the main interfaces between them.
Each of the platform subsystems is described below.
2.2.1. OBC
The on-board computer and data-handling subsystem is based on a distributed control
architecture, whereby each subsystem is governed by one PIC16F877 microcontroller. This PIC
was chosen for Ten-Koh due to its Shinen-2 legacy, which increased the confidence that the controller
would function in-orbit. Using the same microcontroller in all subsystems allowed complete re-use
of the same circuits or even PCB layouts, which drastically reduced the required development time.
Software was also re-used to the maximum extent possible by implementing generic libraries that
could be used across many subsystems. In this way, only the top-level application software layer had
to be developed for every subsystem microcontroller once the libraries were in-place.
The OBC PIC or main controller unit (MCU) is the master of the satellite-wide I2C data bus,
on which all the subsystem microcontrollers are slaves. During the testing of the I2C master–slave
communication architecture with several PIC microcontrollers, we found that when any of the slaves
was powered off, the I2C communication stopped. The cause of this issue was a drop in the I2C
bus voltage, which reduced from the nominal 5 V TTL (transistor–transistor logic) level to around
2 V. By including isolation between the I2C devices that are powered from different EPS power lines,
the I2C data communication bus could continue to operate even if any of the slaves was powered
off. The isolation was implemented with the ADUM1250 IC, which relies on the same magnetic
isolation technology as the ADUM14xx family that has been proposed for small-satellite data handling
application [
]. Moreover, because of the isolation, the I2C bus should remain usable even in the
case of failure of one or more of the microcontrollers. Failure of an isolation IC itself could, in theory,
disable the entire bus. However, due to the use of a flight-proven isolation technology, such risk was
minimized. Another SPI data bus is implemented as backup to enable satellite control in case of I2C
Aerospace 2019,6, 108 5 of 38
bus failures. Due to the limited number of general purpose input/output pins (GPIO) available on the
OBC PIC, only the vital subsystems are connected to this backup SPI bus, namely EPS and COMM.
In case the I2C bus failed, the satellite would still be controllable from ground and operators would
take actions to try and recover full functionality of the spacecraft, e.g., by resetting specific subsystems
or the entire satellite. SPI isolation between all the PICs is implemented with HCPL-0631 optocouplers.
(a) Ten-Koh system architecture.
) Ten-Koh payload architecture including: experiment
control unit (ECU), material mission (MM), charged particle
detector (CPD), and double Langmuir probe (DLP).
Figure 2.
Block diagrams of the Ten-Koh system as well as its payloads. Dashed lines show the limit
of each subsystem and the solid lines indicate the type of interface. Ten-Koh satellite uses I2C as the
main data bus (blue), and SPI (gray) as a backup data bus and direct interface with sensors inside
every subsystem. The 5 V power lines are shown in yellow and the 12 V lines in magenta. The ADS
subsystem was included inside the payload subsystems (PL) for convenience in the physical location
inside the satellite, hence it appears in the bottom block diagram.
The OBC PIC implements the main satellite state-machine software, which reacts to ground
commands and changing conditions with pre-programmed actions. Other subsystem PICs are only
responsible for gathering telemetry from their subsystems and executing OBC commands, with the
exception of the EPS and COMM microcontrollers. These two can either activate battery heaters or
reset the entire satellite, if such a command is received from the ground.
The OBC PIC constantly polls all the currently active subsystem microcontrollers in a sequence
to monitor their status. If any of the subsystems reports a change of state, e.g., a failure, a received
command uplink in case of COMM, or completion of a given task, the OBC PIC responds accordingly.
On-board time, kept by a DS1340 real-time clock (RTC), is also monitored by the OBC such that it can
execute time-tagged commands sent by the ground control.
Analog-to-digital conversion (ADC) on board is performed either by the PIC16F877 in-built ADC,
or when more signals need to be sampled, with an eight-channel, 12-bit AD7927 ADC that uses the
ADR421 voltage reference. The only exception to this are specific payload instruments, which require
more accurate conversion. Data are stored on SD cards, one per subsystem. After reception of an
appropriate telecommand, the OBC will transfer the data from the specified SD card into a buffer,
which will then be downlinked to the ground. In case 5 V logic level of the PICs needs to be shifted
Aerospace 2019,6, 108 6 of 38
to 3.3 V, as required, e.g., by the SD card or the RTC, TXB0104PWR and TXS0102DCTR level-shifters
are used.
2.2.2. COMMS
Telecommunications subsystem includes two hot-redundant uplink chains and two
cold-redundant downlink chains, with one uplink and one dowlink chain governed by one
microcontroller (communications control unit, CCU). The amateur frequency transceivers with flight
heritage were supplied by a Japanese manufacturer ( These are almost
completely self-contained RF circuits and only require an external modem, which was implemented
with the FX614 IC.
Both COMM PICs constantly monitor their respective receivers and notify the OBC in case a valid
uplink command has been received. The OBC then decides which downlink chain to use in order to
send an acknowledgement to the ground-station. Both receivers operate at different frequencies such
that the ground control can choose which uplink chain will receive the uplink, and each command
specifies which downlink chain the OBC should use. This solution allows flexible in-flight debugging
of the system.
2.2.3. ADS
The attitude determination subsystem philosophy is outlined in Figure 3. Raw measurements
from coarse sun-sensors, gyroscopes and magnetometer are processed on-ground in order to estimate
the attitude. Because attitude control has not been implemented, attitude determination on-board was
not required, which further simplified the on-board software.
The 12 coarse sun-sensors were implemented using S4349 photodiodes sampled with the AD7927
ADC, with one diode placed on each solar panel of the satellite. Triple modular hot-redundant
A3G4250d gyroscopes were used to provide attitude-rate measurements, while the magnetic field
is sensed by the HMC2003 three-axis magnetometer and two 16-bit ADCs. All the ADS sensors are
sequentially sampled over a SPI data bus by one PIC microcontroller in order to ensure that all the
readings are synchronized, which is required for attitude determination. A detailed description of the
ADS subsystem is included in [17].
While the spacecraft is in view of the ground station, the ADS sensor data can be downlinked in
real-time and combined with downlink signal measurements done by the ground station. This enables
higher attitude determination accuracy during ground station passes.
Figure 3.
Attitude determination philosophy for the Ten-Koh satellite. The magnetometer instrument
is used for space environment purposes as well as for ADS.
Aerospace 2019,6, 108 7 of 38
2.2.4. EPS
The main power source of Ten-Koh consists of 12 solar panels composed of five triple-junction
solar cells each, with two SPV1040 MPPT (maximum power point tracking) circuits per solar panel:
one for managing two cells and one for managing the remaining three cells. All 12 solar panels are
connected in parallel in an unregulated power bus that is used to charge two hot-redundant batteries.
Each of the two batteries is composed of four individual Li-ion cells connected in parallel. The complete
outputs of both batteries are also connected in parallel in order to form a battery-regulated bus.
The battery-regulated power bus is an input into the LT1370 voltage regulators, which generate 5
and 12 V regulated buses. These buses then feed over-current protection (OCP) circuits implemented
with LTC4361 and LT4363 ICs, which in turn deliver power at regulated voltages to the different
subsystems. All subsystems were designed to use these voltages to reduce the complexity of the entire
system. In certain cases, however, 3.3 V levels are required. These are obtained from the 5 V lines with
LD1117 linear regulators within the subsystems.
The different power lines per subsystem are:
Two independent 5 V lines for the payloads—one for the experiment control unit, ECU and one
for the double Langmuir probe, DLP.
One 12 V line for the charged particle detector (CPD) instrument.
One 12 V line for the magnetometer instrument.
Two 5 V lines, one for each of the redundant communications subsystems.
One 5 V line for the attitude determination subsystem (ADS) and all solar panels telemetry
acquisition circuits.
One 5 V line for the entire EPS subsystem.
One 5 V line for the OBC subsystem.
Each power line can be controlled (switched on and off) independently from the EPS controller
PIC. The default state of the power lines is encoded in hardware to make sure that the OBC and COMM
subsystems are on even in case of failure of the EPS controller. This PIC also gathers voltage, current
and temperature telemetries from the entire subsystem.
The total power needed to run all the platform subsystems (OBC, EPS, and COMMS) is 3.3 W.
When the payloads are operated, this value increases to at most 9.5 W, which is the case of running
the platform plus the CPD and DLP (worst case allowable for payload operation). Depending on the
experiment being conducted, the power consumption can be reduced to:
Material mission (MM): 4.8 W.
Charged Particle Detector (CPD): 8.8 W.
Double Langmuir Probe (DLP): 5.5 W.
Attitude Determination Subsystem (ADS): 5.5 W.
Thermal Switch (TSW): 5.5 W.
Ultracapacitor (UCP): 5 W.
From the power generation perspective, the minimum available power is 11 W, which leaves
sufficient margin to recharge the battery during sunlit parts of the orbit. A detailed description of the
solar panels designed for Ten-Koh is shown in [18].
2.3. Payload
The main payload consists of five individual instruments, the main characteristics of which
are summarized in Table 1. One of the main payloads is the Charged Particle Detector (CPD),
which was developed by Prairie View A&M University, TX, USA and the Space Research and
Technology Institute—Bulgarian Academy of Sciences, Bulgaria. This system consists of eight CMOS
Aerospace 2019,6, 108 8 of 38
(Complementary Metal Oxide Semiconductor) detectors mounted on five faces of a cube, and the
Bulgarian Liulin-type detector [
] mounted on the top of the complete assembly. The CPD instrument
contributes to all four primary mission objectives.
Another payload of Ten-Koh is a set of Double Langmuir Probes (DLP). This experiment is
intended to study the plasma and the sheath formation around the satellite as it is spinning along its
orbit. Satellite charging, which can occur at surface and internal level, is an important subject of study
because it is related to satellite anomalies and is directly impacted by the space weather conditions.
Satellite charging depends on the interaction of the spacecraft materials, their thickness, and charged
particles energy. Fluxes of electrons ranging from 10 to 100 keV can produce surface charging,
while electrons with energy bigger than 100 keV can produce internal charging (deep dielectric
charging) [
]. The plasma flow around a satellite in LEO exhibits a behavior similar to the flow of
a neutral gas: there is a region of compression on the side where the plasma flow impinges upon the
moving spacecraft, and a wake region behind the body of the satellite (see Figure 4a). For satellites
that have no attitude control, the modeling of the plasma interaction with the satellite body becomes
a complex task. Therefore, in-situ measurements provided by the described experiment would be
a valuable source of information useful for the validation of the plasma models.
The DLP system is composed of two spherical, 10
m gold-plated electrodes, as shown in
Figure 5a, which are placed outside of the spacecraft structure. The electrodes are mounted on an
isolating glass fiber reinforced polymer (GFRP) plate to avoid any electrical contact with the CFRP
structure, thus keep the electrodes floating with respect to the spacecraft ground. The shafts supporting
the spheres are coated with alumina to ensure that only the spheres attract plasma particles, thus
making the interpretation of results easier. The DLP control circuit, housed inside the spacecraft, uses
a 16-bit digital-to-analog converter (DAC) to generate the DC biasing voltage. This allows the sweep
duration to be tailored such that the plasma in different locations around the spinning spacecraft can
be characterized, as shown in Figures 4a,b and 5a,b. The current flowing between the probes through
the plasma and the biasing voltage are measured with a AD7927 12-bit ADC. The measurement
circuit is isolated from the common ground of the spacecraft, which enables the plasma parameters to
be measured.
The purpose of the material mission (MM) is to expose three different samples of new material
for space applications developed by Okuyama laboratory of Kyutech made of carbon fiber (CF) and
PEEK resin:
Sample 1: CF/PEEK with no coating.
Sample 2: CF/PEEK with a special coating (silsesquioxane) to protect against atomic oxygen.
Sample 3: CF/PEEK with a special coating (yttrium oxide) to protect against UV.
The system measures different parameters, including temperature of and strain inside the samples
in order to quantify the changes in the coefficient of thermal expansion (CTE) of the new materials.
By measuring the changes in CTE, internal micromechanical modifications in the chemical composition
of the samples caused by the effects of space environment can be discerned. Testing in this way allows
assessment of material survivability and reliability for use in space. The temperatures are sensed with
AD590 temperature transducers sampled by the AD7927 ADC. The ADS1220, a 24-bit ADC, and COTS
strain-gauges are used to measure strain in two orthogonal directions inside the samples.
Aerospace 2019,6, 108 9 of 38
(a) (b)
(c) (d)
Figure 4.
Representation of the DLP measuring system: (
) I–V (current–voltage) sweep locations
around a spinning spacecraft; and (
) 1-D model representing the electrodes inside and outside of
the plasma sheath.
The payloads are controlled using an experiment control unit (ECU) PIC. The MM constantly
measures the samples’ properties and, therefore, implements a dedicated controller not to block
or interrupt the CPD and DLP operations. The magnetometer was already described in the scope
of the ADS subsystem because it serves a dual purpose of attitude determination and scientific
data generation.
Another secondary payload consist of a GNSS receiver, which was included to fulfill two purposes:
1. To increase the spatial and temporal localization of other payload measurements.
To investigate the GNSS signal disruption, total or partial, due to ionospheric varying conditions.
The FireFly GNSS receiver with an active antenna model TW2406 (for GPS/GLONASS L1 band)
were selected. This GNSS receiver was selected due to its low size (17.0 by 22.4 mm) and low power
consumption (150 mW). It is considered one of the world’s smallest space-capable GNSS receivers
according to the manufacturer’s information (
pdf). This receiver has been used on-board small rockets launched by JAXA and other university
satellites such as the Fireant Cubesat [
]. The receiver was included within the same PCB as the
MM to optimize the space available in the top of the spacecraft, and to minimize the number of PIC
microcontrollers. However, due to the lack of resources in the MM PIC microcontroller to handle both
the MM experiment and the GNSS readings, and due to the low priority assigned to this secondary
payload, the GNSS receiver has not been used. Therefore, it is not mentioned any further in this paper.
Aerospace 2019,6, 108 10 of 38
Figure 5.
Representation of the DLP measuring system: (
) DLP electrode made of solid aluminum,
coated with 10
m of gold in the spherical tip, ceramic material (in white color) in the shaft for
isolation and FR4 plate (in gray) on the base for isolation purposes from the structure of the spacecraft;
and (
) the DLP measuring circuit block diagram. The dashed line indicates the DLP measuring circuit
physically implemented in a single PCB.
Aerospace 2019,6, 108 11 of 38
Table 1. Main payload instruments.
Item Instrument/Sensor Measurement Range
Charged particle detector
based on CMOS sensors
6 detectors for protons,
electrons and ions.
2 x-ray detectors
The final range will be defined based on
ground calibration and orbital data.
Dosimeter for space radiation.
Liulin type dosimeter
Dose rate [Gy/h]: 1.12 ×108to 0.188
Flux [particles/cm2/s]: 0.01–10,000
Deposited Energy [MeV]: 0.081–20.833
Material mission Strain measurements
Temperature measurements
Strain: 10,000 ×106to 1 ×106
Temperature [C]: 55–150
Current [A]: +/10 ×106
Minimum current [A]: 80 ×1012
Voltage [V]: +/11
Magnetic field measurements
Magnetometer based on
3-axis magnetoresistive
2.4. Ten-Koh General Arrangement
Figure 6shows the arrangement of the subsystem assemblies in the Ten-Koh structure. Most of
the subsystems parts, including the payload instruments control electronics and several instruments,
are mounted on the internal structure.
The material mission (MM) payload control electronics circuits are located inside the structure,
while the material samples are placed directly above and exposed to the space environment.
Similarly, the double Langmuir probe (DLP) electrodes are located on the outside of the external
structure of the satellite. However, its control electronics circuits are housed in the internal structure
shielded in a 3 mm-thick aluminum box.
The magnetometer instrument is mounted on the top panel in an aluminum box. The selection of
the location and materials for the mechanical interface of the magnetometer instrument is discussed
below as part of the challenges in the design of the Ten-Koh satellite.
2.5. Satellite Concept of Operations
For all experiments and mission objectives, the concept of operations is shown in Figure 7. There is
one main ground station (MGS) located at Kyutech, from where all missions are planned and executed,
and housekeeping and payload data are downlinked. The mission was defined to start operations in a
period of 2–3 weeks after the launch by following the next sequence of events:
Launch on 29 October 2018 at 13:10 Japan Standard Time (04:10 UTC) from the Tanegashima
Space Center in Japan, onboard the H-2A rocket number F40 as a piggyback payload of the
GOSAT-2 satellite.
Separation from the rocket 2000 s after liftoff at an altitude of 623 km and with an orbital velocity
of 7.5 km/s. The preliminary orbit was estimated to plan tracking of the satellite during the first
hours after separation, until a more precise ephemeris was obtained later: semi-major axis of
6969.772 km; eccentricity of 0.002586; inclination of 97.835 deg; argument of perigee of 27.546 deg;
longitude of the ascending node of 313.285 deg; and true anomaly of 212.990 deg. All data
were provided by the Japan Aerospace Exploration Agency (JAXA) prior to the launch [
Nominal operations were carried out using Two Line Element sets (TLEs), as customary for
small satellites.
After the separation and orbit insertion, and once all systems onboard were verified, the mission
operations started on 19 November 2018.
During each communication session with the main ground station located in Kyutech,
housekeeping telemetry and mission data are downlinked and missions are executed according
to a previous plan (see the following process of mission planning and execution in Steps A–D).
Aerospace 2019,6, 108 12 of 38
Figure 6.
The payload arrangement onboard the spacecraft: external arrangement (
internal arrangement (b,c); and internal structure dimensions in millimeters (d).
Based on the above key events, the mission operation phases given in Table 2have been defined.
Each phase refers to a section of the life cycle of Ten-Koh satellite, which itself starts with the launch
and lasts until the decommissioning of the mission.
The process of performing all missions on-board Ten-Koh (Processes 3 and 4 in the concept of
operations above) is as follows:
Aerospace 2019,6, 108 13 of 38
An experiment is planned to be executed in a selected position and epoch along the orbit with a
specified time duration.
A set of uplink commands is configured and scheduled for commanding the satellite in a selected
pass over the MGS in Japan.
During the designated pass for operations, the satellite health status is confirmed first by
housekeeping telemetry. Then, the uplink command is received by the satellite and the OBC
configures the payload instruments and the required subsystems to provide the resources for the
mission to be executed.
The mission experiment is executed and the housekeeping and missions data are stored in
different memories on-board the satellite to be downlinked during the next passes over the MGS.
The process in Steps A–D is repeated for every new mission experiment.
Table 2. Satellite mission operation phases.
Mission Phase Duration Description
LEOP: Launch
early operations
2–3 weeks
Starts after separation from the
launch vehicle.
Subsystems and instruments are
powered on.
Comprehensive satellite
subsystem checkout.
Recording of Satellite Telemetry
Data (STD).
Downloading and analysis of
scientific and technical data by
ground control station.
Checking the operation of payload
instruments (by real time mode
Transition to normal operation.
1 year starting after
3 weeks from the
Science data collection.
Long and short term planning of
payload instruments.
Communication sessions with
ground stations in charge of
controlling the satellite.
1 year, starting after
Normal Operation
Variations in science operations.
Orbit decay monitoring.
4 End of mission >3 years Planning and execution of the end
of life operation.
Aerospace 2019,6, 108 14 of 38
Figure 7.
Satellite concept of operations as explained in sequence from 1 to 4. In Processes 3 and 4,
a mission is configured to be executed and later payload and housekeeping data are downlinked. This
mission configuration–execution–downlinking occurs following the execution of Steps A–D. After
launch and when the nominal operations started, only Processes 3 and 4 are present as part of the
nominal operation.
2.6. Design of Mission Operations
The Japanese Aerospace Agency (JAXA) offered free launch slots for the GOSAT-2 launch in
early 2017, and the Ten-Koh mission design was based on general estimates and assumptions for
a Sun-synchronous polar orbit with an altitude of 613 km and 97.8 degrees inclination (the final orbit
parameters were not confirmed until a few months prior to the launch). The previously described
experiments and the associated six instruments were conceptualized during a brainstorming process
and an official proposal was submitted to the JAXA as a bid for the GOSAT-2 launch opportunity,
which was offered to Japanese universities. After the launch, the final orbit was confirmed to have the
following parameters:
Perigee: 593 km.
Apogee: 615 km.
Inclination: 97.8 degree (retrograde orbit).
Semimajor axis: 6975 km.
Orbital period: 96.6 min.
Eclipse duration: 47.7 min.
Number of revisit times per day over Kyutech (Japan): 4.
Revisit cycle: 9 days.
Local equator crossing time of descending node: 13:00 h.
Even though Ten-Koh can pass over the Kyutech ground station in Japan up to four times per day
(twice at around 13:00 and twice at 01:00), only passes with more than 15 degrees of elevation are used.
This is done in order to establish reliable communication between the ground station and the satellite
(i.e., to avoid the loss of data packets). Ultimately, an average of two passes per day are usable.
Satellite Operation Modes
The operation of the satellite is divided in the following operation modes. These are linked to the
specific events from the concept of operations previously explained in Figure 7via the number shown
in parenthesis:
LEOP mode (1).
Communication sessions mode (4).
Aerospace 2019,6, 108 15 of 38
Telemetry mode (4).
Real-time mode (4).
Normal mode (2).
Payload operation mode (3, 4).
Satellite log mode (4).
Direct command mode (4).
Set SD card saving data address mode (4).
During the payload mode, the MGS sends a command or a series of commands for executing
payload operation according to the following mission experiments:
CPD: measurements of CPD are sampled and stored it in the ECU’s SD card.
DLP: measurements of DLP are sampled and stored it in the ECU’s SD card.
MM: measurements of MM are sampled and stored it in the MM’s SD card.
ADS: measurements of ADS (magnetometer, gyroscopes and Sun sensors) are sampled and stored
it in the ADS controller’s SD card.
UCP experiment: measurements of UCP are sampled and stored it in the UCP’s SD card.
Thermal switch: measurements of the thermal switch are sampled and stored it in the UCP’s
SD card.
Depending on the payload instrument to be operated, the configuration for the operation must
be included in the uplink command for the experiment to run. For the cases of CPD, DLP and ADS,
the payload mode uses a definition of two operating windows along the orbit, which allows for the
CPD and DLP cases to sample data from each instrument in two different locations of the satellite orbit.
For instance, the two missions can be performed over the north and south poles, or in two opposite
regions over the equator, as shown in Figure 8.
2.7. Challenges Faced during the Project
The Ten-Koh satellite was developed to meet the mission objectives specified in Section 2.1 while
meeting constraints such as the size and mass limits, the orbit driven by the launch’s primary payloads,
the schedule dictated by JAXA, and the available resources and small team size. Especially the schedule
and resource constraints have driven some of the key decisions throughout the project, both technical
and programmatic in nature. The main challenges and solutions to them are summarized in Table 3
and then described individually in more depth.
2.7.1. Available Generated Power
The minimum available power in the worst-case scenario was computed to be 11 W, which was
found sufficient for the maximum expected power consumption of 9.5 W. The EPS and the batteries
were sized to be able to support such power consumption in eclipse for up to 30 min, which was
enough to gather CPD data around the poles, regions of specific interest due to abundance of charged
particles. The EPS was not sized to be able to support maximum power consumption indefinitely
due to the limited surface area that could be covered with solar cells. Using a deployable solar array
could remedy this area constraint but would require development of complex on-board software and
mechanisms, and was thus deemed unfeasible. Attitude control would also be needed if a deployable
array was used, which is described in a separate section.
Aerospace 2019,6, 108 16 of 38
Table 3. Design challenges and adopted solutions.
Challenge and Its Origin Solution
aChallenge: Available generated power
Origin: SmallSat constraint
Maximized solar array area + payload
operations with reduced time duration.
Challenge: Available space location,
mass and structure interface for some
payload instruments
Origin: SmallSat constraint
Platform subsystems were compactly
accommodated at the bottom of the central
internal structure and the inside walls of the
external structure to maximize payload space.
Structure effects on the measured data need to
be corrected for during result interpretation.
Challenge: Amount of data generated by the
mission instruments
Origin: SmallSat constraint
On-board storage capacity of up to 2 GB for
the main payload instruments, EPS and ADS
subsystems each.
dChallenge: Structure design heritage
Origin: Schedule for delivery on time
Shinen-2’s legacy structure with location for
Ten-Koh components, considering thermal
control, fields of view of antennas and sensors.
Challenge: Attitude determination and
control vs attitude determination
Origin: Schedule for delivery on time
and keeping the CPD and magnetometer
readings with reduced influence
An attitude determination only approach was
adopted to meet schedule and cost constraints.
Challenge: COTS design heritage from
previous missions
Origin: Reliability and budget constraints
Use of avionics components from Shinen-2
and other small satellite missions wherever
Challenge: Magnetic measurement system
with no deployable boom
Origin: Schedule constraint
Magnetometer instrument location was
designated in the top panel of the satellite
to reduce complexity from boom deployable
system while keeping the schedule and
measurements quality.
hChallenge: Launch date fixed by JAXA
Origin: Schedule constraint
Adopted a modular design approach,
with re-use of the same microcontroller
and other circuitry in all the subsystems. This
also maximized software reuse.
2.7.2. Available Space Location, Mass and Structure Interface for Some Payload Instruments
The payload instruments were located in the upper half of the satellite in order to provide
an unobstructed view for the CPD particle detectors, expose the material samples to the space
environment, and allow CPD and DLP to be controlled by the same ECU, thus allowing synchronized
measurements. Initially, controlling the GNSS receiver and the magnetometer from the same
microcontroller was foreseen but was eventually found unfeasible due to PIC16F877 hardware
limitations. However, these payloads were kept in their original locations to avoid large configuration
changes late-on in the project. Only the microcontroller responsible for these two instruments
was changed.
Aerospace 2019,6, 108 17 of 38
Figure 8.
Operation of the main payloads as described in Processes 3 and 4 in Section 2.5. (
Communication sessions between the satellite and Kyutech ground station in Japan occur in two
periods, during the day at around 13:00 and during night time at around 01:00. (
) Mission execution
for CPD, DLP and ADS in a specified region that is pre-selected and configured via a command during
the communication sessions with the satellite.
In the case of the satellite’s top panel, all the space available is dedicated to accommodate
the material mission assembly, the magnetometer instrument and a GNSS receiver antenna. The
magnetometer was placed on the top panel, as far as possible from other subsystems and particularly
the switching voltage regulators in order to reduce the noise in the magnetic field measurements
(explained further in section h).
The DLP electrodes were located on opposite sides of one of the hexagonal panels close to the
ECU, with their exact location and length tailored so as to comply with the allowed physical envelope
Aerospace 2019,6, 108 18 of 38
of 50 cm
50 cm
50 cm dictated by the HII-A rocket. Due to this physical envelope limit and in
order to avoid time-consuming development of an expandable electrode system, the DLP electrodes
have fixed length. However, when fixing the electrodes close to the satellite structure, the measurement
system must be able to measure plasma properties when the electrodes are located inside and outside
of the plasma sheath, which implies the ability to measure exceedingly low current values. The
IRI-2016 model [
] at the time of Ten-Koh launch (2018) was employed to find the expected plasma
density in the range from 1.0409
to 2.0248
(mean 7.0153
), and an
electron temperature of 2407.8 K (0.2 eV). The resulting Debye length of 1.28 cm suggested that the
chosen probe length would indeed locate the probes inside and outside of the sheath, depending
on spacecraft attitude (schematically shown in Figure 4). Thus, the size of the probes was found
appropriate to address primary Primary Objective 1 and the measurement circuit was designed and
tested accordingly, as shown in Section 3.1.
2.7.3. Amount of Data Generated by the Mission Instruments
Ten-Koh has been equipped with four SD cards of 2 GB each in order to provide on-board
storage for the mission, ADS, and housekeeping data. The DLP instrument generates the largest
amount of data per single measurement with 3000 bytes in the high-resolution mode, followed by
the CPD with 2528 bytes (2000 from CMOS detectors and 528 from the Liulin detector). Table 4
shows the main mission instruments listed in descending order for the amount of data generated
considering the different mission scenarios. Storing data on-board presents no difficulties due to the
use of SD cards. However, the downlink of relatively large amounts of data can become problematic
when using only the amateur UHF frequencies at the rate of 9.6 kbps supported by the used COTS
transmitter. Thus, the experiment duration has to be tailored depending on team priorities and
operational conditions. Increasing the number of downlink locations would improve the system
data throughput at the expense of more energy needed per-orbit to operate the transmitter for longer.
Ultimately, this approach was not adopted due to insufficient time to prepare additional ground
stations before the launch. During the first three months of operations, maximum duration of 10 min
was planned for each of the main missions. Depending on the real downlink capability, extending the
duration for the main missions would be considered.
Table 4.
Data generation for the main mission instruments as a function of the duration of each
experiment for one ground station access case.
Mission Instrument
Amount of Data
Generated Per Single
Measurement Per
Channel [Bytes/ch]
Maximum Time
Duration of the
Experiment [minutes]
Data Generated for the
Maximum Time Duration
for 2 Channels [Bytes]
3000 bytes (hi-res)
1500 bytes (med-res)
600 bytes (low-res)
15 min
900,000 bytes (hi-res)
672,000 bytes (med-res)
180,000 bytes (low-res)
CPD 2528 bytes 15 min 75,840 bytes
ADS sensors 64 bytes 15 min 57,600 bytes
Material mission 60 bytes 5 min 18,000 bytes
2.7.4. Structure Design Heritage
Using a heritage Shinen-2 structure had the advantage of reusing the same composite mold and
thus lowering the total development cost. In addition, it increased the confidence in the ability of
the structure to withstand the launch environment. However, the arrangement of the subsystems
in the structure has been changed completely in order to reduce the amount of harness as well as
to provide the required space and viewing conditions for the payloads. The location of the DLP
electrodes is described in Section 2.3, Figure 4, and Figure 6c. Besides that, the fields of view of the
Aerospace 2019,6, 108 19 of 38
CMOS sensors were kept as unobstructed as possible by placing the CPD in the center of the internal
structure, sufficiently below the upper aluminum boxes of the ECU, DLP electronics and ADS.
Both Shinen-2 and Ten-Koh structures share the same shape and size but the material is different:
Ten-Koh uses CFRP (carbon fiber + epoxy resin), while the external structure of Shinen-2 was based
on CFRTP/PEEK material. CFRP/epoxy composite materials have been tested for their use in
space, where the atomic oxygen and UV radiation represent the biggest challenge for graphite
epoxy composites [
]. In this regard, CFRP/epoxy has been shown to have increased durability in
LEO compared with CFRTP/PEEK. Both CFRTP and CFRP have similar radiation permeability [
and thus would both provide similar shielding to the electronics and the CPD detectors. During the
structural-thermal model (STM) tests, it was confirmed that the spacecraft can be qualified for the
vibration environment of H-IIA (more than 100 Hz in the longitudinal direction, with
6.0 G
compression and +5.0 G tensile, and more than 50 Hz in the lateral direction, with
6.0 G of
compression and tensile).
2.7.5. Attitude Determination and Control vs. Attitude Determination
Performing only attitude determination without attitude control was chosen in order to reduce
the satellite development time. Only passive magnetic attitude control was initially deemed feasible
for schedule reasons and was proposed early in the design process. However, it would degrade the
CPD and magnetometer readings and thus was discarded. From the point of view of the mission
objectives, being able to understand the direction in which the particle/plasma measurements were
taken was sufficient.
2.7.6. COTS Design Heritage From Previous Missions
Ten-Koh avionics are based on Shinen-2’s COTS components and other Kyutech satellite missions
that have been flight-proven in space. However, in certain cases, no heritage components from
previous Kyutech missions were found. In these instances, literature review was conducted to find
candidate COTS components. ADUM1250 and ADUM14xx families of magnetic isolation devices are
of note in this respect because the entire satellite data handling subsystem relies on these ICs, and
thus choosing components likely to survive space environment was needed [
]. The ADUM1250 IC
is used for I2C isolation purposes while the ADUM14xx family is used for SPI and general-purpose
digital signals with 5V-TTL compatible levels. If no components with reported flight heritage were
found, device packaging, used technology, and temperature range were considered in order to select
the IC from amongst the commercially-available ones.
2.7.7. Magnetic Measurement System With No Deployable Boom
The magnetometer instrument was installed on the top structural panel of the satellite next to the
material samples (see Figure 6). Calibration was performed in order to validate whether the instrument
was able to perform magnetic field measurements inside the expected range along the satellite orbit,
and that it was not influenced beyond acceptable levels by the spacecraft self-induced magnetic noise.
A sensitive and linear output inside the expected range was obtained, while confirming that the
induced noise (offset from the spacecraft) was an order of magnitude below the instrument’s range
limits of
T (
nT). Because the different offset values were constant (see Figure 9),
it was possible to characterize the different values produced by the different modes of operation of the
satellite, thus enabling the removal of magnetic noise from the measurements during post-processing
of the data. This calibration and post-processing will be described in detail in a dedicated publication
because it requires extensive explanation that cannot be accommodated in this paper.
Aerospace 2019,6, 108 20 of 38
Figure 9.
Magnetometer calibration tests. Offset imposed by the satellite to each axis when:
no electronics are ON, and under specific subsystems operating conditions. The results will be
used to remove the offset to the magnetometer measurements when the satellite is under specific
operating conditions.
2.7.8. Launch Date Fixed by JAXA
The satellite launch slot was awarded to the team through a competition hosted by JAXA.
However, to be eligible for the launch, the satellite had to be delivered and qualified for flight
on-time with no possibility of postponing the launch, which would take place in case of primary
launch payloads.
To meet the rigorous schedule, the system was kept as simple as possible: for example, no
deployable boom for the magnetometer and no attitude control were used, as described above.
Moreover, a modular design approach was adopted, where the same microcontroller and other
circuitry were reused in all the subsystems. This also maximized the possibility to re-use the
same lower-level software across the subsystems. Lastly, a proto-flight approach was adopted for
system-level thermal-vacuum tests in order not to lock away the engineering qualification model in a
vacuum chamber at the peak of the development activities.
3. Results
3.1. Ground Tests
The main payload instruments were tested and calibrated to verify that the experiments can
provide measurements inside the defined range. For the MM, DLP and magnetometer, the testing and
calibration were carried out using the facilities of Kyushu Institute of Technology in Japan. For the CPD
case, which includes the CMOS and Liulin detectors, the tests were conducted at Prairie View A&M
University and the Space Research and Technology Institute of the Bulgarian Academy of Sciences,
respectively. Details about the Liulin detector calibration can be found in [
], while the CMOS
detectors still need an extensive calibration that is currently underway at Prairie View A&M University.
Double Langmuir Probe (DLP)
Figure 10a shows the ground test results of the DLP plasma measurement system. The different
I–V curves for each case correspond to varying numbers of samples, and hence the speed at which
the applied potential was being changed. This resulted in different coupling of the circuit with the RF
Aerospace 2019,6, 108 21 of 38
plasma source in the chamber. For a large number of samples, the DLP generates smaller voltage steps
(step size). Considering that the sheath formation varies as a function of the biasing voltage applied to
the probes (VB) with respect to the floating potential Vf (zero net current or equal flow of electrons and
ions), by setting the probes potential above the Vf, a net positive current due to plasma electrons flows
into the plasma. If VB is set below Vf, then a net negative current of ions flows out from the plasma.
By biasing the probes with negative and positive potentials, the typical I–V curve is obtained (see left
of Figure 10a). After different tests performed at the LEO plasma chamber located in the facilities of
the Laboratory of Spacecraft Environment INteraction Engineering (LaSEINE) in Kyutech, the
to +10 V biasing range was deducted as optimal to obtain a full I–V curve in orbit. This range allows
computing the electron density and temperature of the plasma surrounding the spacecraft according
to the tests and estimated parameters from the IRI 2016 model [24] for the referred epoch.
When biasing the voltage across the probes in smaller steps (more samples), the DLP system takes
longer time to sweep the voltage from
10 V to 10 V. When selecting a bigger step size (fewer samples),
the DLP sweeps the voltage across the probes faster. This can be seen as a change in the frequency of
the biasing voltage. Because the plasma source employed in the testing was of RF origin, the plasma
potential fluctuated at the RF frequency and its harmonics. The overall effect appears as a wider I–V
curve that creates a higher value of the electron temperature and shifts the floating potential to a more
negative value, as described in [
]. To reduce the effect of the RF coupling with the measuring circuit,
the selected step size used during the testing was between 500 and 1000, which corresponds to 65 and
131 samples. This was validated with the use of a source-meter instrument with an RF compensation
stage and requesting the same number of samples.
The final ground test for the DLP instrument development was the calibration of the system.
The fastest way to perform calibration of the measuring unit was by the use of a relative calibration
with the same SMU2400 instrument. By comparing resistance measurements from current and voltage
readings, it was possible to obtain the sensitivity and linearity of the DLP system relative to the
SMU2400 and the resistor value. For this purpose, a resistor of 1 G
ohms), and 1% precision
value, was connected between the two probes at room temperature. Then, the voltage was biased from
10 V to 10 V across the resistor. During this test, no plasma environment conditions were applied.
Figure 10b,c shows the results from the calibration of the DLP instrument while Table 5includes the
calibration results for the four channels in the DLP system.
Results from the ground tests in the LEO plasma chamber show that the Ten-Koh DLP is provided
with an electron temperature of 1.38 eV and a plasma density of 1.91
. As a reference,
the LEO plasma chamber produces plasma densities in the range of 10
, meaning that the
DLP results were plausible.
Table 5. Calibration results for all four channels of the Ten-Koh DLP instrument.
Value [G]
Relative Error from
1 [G] Resistor
Nominal Value in %
Relative Error from
SMU Measured
Resistance in %
1 (CHA1) 1.0010 ×1090.1 0.82
2 (CHB1) 1.0180 ×1091.8 2.53
3 (CHA2) 1.048 ×1094.8 5.5
4 (CHB2) 9.944 ×1080.56 0.15
SMU2400 9.929 ×1080.71 ——
Aerospace 2019,6, 108 22 of 38
Figure 10.
Ground tests results for: (
) DLP I–V curve results for different number of samples (DAC
step size) and the results comparison to an ideal DLP I–V curve on the left; (
) DLP measuring circuit
comparison with the SMU for the same plasma conditions inside the LEO plasma chamber; and (
DLP sensitivity and linearity relative calibration by using a 1 G
resistor and the SMU2400 instrument.
The DLP system has a linear response as expected inside its operating range. The current saturation in
channel CHA2 is due to its higher gain compared with the other channels.
Aerospace 2019,6, 108 23 of 38
During the ground test of the DLP system, the EM satellite structure and the DLP electrodes
mounted on it, as well as the plasma source were kept static inside the plasma chamber. The face of
the satellite structure where the electrodes were mounted was positioned toward the direction of the
plasma source (0 deg) inside the chamber and data were acquired. Then, after removal of the plasma
and vacuum, the satellite structure was rotated by 180 degrees, i.e., the electrodes were facing away
from the plasma source. Another set of measurements was acquired in this configuration. Only these
two tests were conducted in order not to occupy the plasma chamber for longer than absolutely
necessary, due to the high demand for the facility. However, the two tested orientations, with the DLP
electrodes facing towards and away from the plasma flow, captured the two extreme thicknesses of the
sheath in which the experiment would operate. This was associated with a reduction in the magnitude
of the measured current in an order of 101A.
Under flight conditions, the lack of attitude control invariably affects the results as the alignment of
the DLP electrodes with respect to plasma flow varies continuously. In this regard, attitude knowledge
is critical to provide the location of each DLP sweep and to understand how quickly the attitude is
changing. The sampling rate also plays an important role because sweeps with higher resolutions
require more time for the data acquisition and, therefore, the attitude will change more during the
sweep. Because the attitude and attitude rates were uncertain, the experiment was designed to be
configurable by software via an up-link command. The parameters that can be set by software are:
Resolution: The number of samples for each I–V sweep can be a set of up to 16 bits voltage
increments. Three exemplar cases are presented in Table 4—different resolutions can be
implemented depending on the satellite rotation conditions.
Biasing voltage range: The biasing voltage range can be selected from
10 to +10 V (default),
or be any other combination in that range. The only constraint is that the biasing voltage always
goes from the most negative value to the most positive one.
Any combination of the mentioned parameters can be set for the DLP experiment to run. With this
flexibility, it is possible to adjust the configuration of the experiment to avoid limiting the sampling rate
and the charging conditions around the electrodes. However, as described in Sections 4.2.2 and 4.2.3,
due to several failures, the DLP experiment could not be operated in-orbit.
3.2. In-Orbit Results
3.2.1. CPD Instrument
Five datasets were obtained from the Liulin detector during the early operations phase of the
Ten-Koh mission. Table 6shows the processed data from three different days when Ten-Koh satellite
was flying over Japan. The Liulin detector was operated in a real-time mode, which means that
the measured data were transmitted to the ground station without storing them on the SD card.
These datasets represent the first in-orbit measurements from the CPD instrument. The real-time mode
of the CPD was selected in order to test the system behavior when turning on the major load on-board.
The data obtained from the early operations show that, for close L-shell [
] values (1.129 to
1.206) and around the same time conditions (afternoon local time), the fluence has an average value of
0.28 particles/cm
/s. For a GEO field with no perturbations and short duration measurements period
(less than 1 min), the particles flux is expected to be in this order, i.e., the CPD data are consistent with
the expectations. However, the dose rate from the CPD data shows fluctuations. The average dose rate
measured was 1.74
Gy/h, which is equivalent to 4.832
rad/s. The low dose rate measured in
the second spectrum on 18 November 2018, comes from only low energy channels registering hits in
the instrument. Figure 11 shows the five spectra and the corresponding average spectrum (in black)
from the real time datasets. The horizontal axis in Figure 11 represents the deposited energy (in 256
channels) for the energy range from 0.081 to 20.833 MeV of the Liulin detector. In yellow, the spectrum
with the lowest dose rate shows that the instrument registered only deposited energy in the low energy
Aerospace 2019,6, 108 24 of 38
channels, which in turn reduced the received dose. Similarly, the second spectrum on 18 November
2018, measured deposited energy in the low region of the instrument’s energy range. In both cases, the
total dose rate resulted in less than 1
Gy/h. The cause of the fluctuations in the dose rate is associated
with the rotation of the satellite, which misaligned the CPD-Liulin detector with respect to the GEO
field lines as it was traveling around the Earth. This is observed for the spectrum 2 on 12 November
and 18 November 2018, where the precedent spectrum in each case measured a higher dose rate than
spectrum number 2. Figure 12 shows the geographic locations of the real-time mode CPD mission
performed on Nov 2018.
Figure 11.
Dose rate and flux spectra obtained from the real time operation mode for the first
CPD-Liulin data measurements on November 2018 during the early operation phase of Ten-Koh
satellite: (
) dose rate spectrum for each dataset and the average spectrum measured inside Ten-Koh
spacecraft; and (
) integral flux spectrum for the 5 real time datasets and the average flux spectrum
measured inside the spacecraft.
Aerospace 2019,6, 108 25 of 38
Table 6.
The Liulin spectra datasets obtained during real-time operations of the instrument. The measured dose rate and flux where obtained from a total exposition
time of 15.286 s for each spectrum dataset. Epoch is in UTC even though the operations were conducted using the Japanese Standard Time (JST).
Spectrum Epoch UTC Geographic Location of the
Dose Rate
[µGy/h] Dose Rate [rad/s] Flux
108 November 2018
Lat [deg]: 25.98
Lon [deg]: 132.67
Altitude [km]: 609.33
1.144 1.93 5.444 ×1080.33
112 November 2018
Lat [deg]: 29.55
Lon [deg]: 123.09
Altitude [km]: 608.41
1.206 2.66 7.389 ×1080.39
212 November 2018
Lat [deg]: 27.16
Lon [deg]: 122.50
Altitude [km]: 607.76
1.161 0.39 1.083x1080.2
118 November 2018
Lat [deg]: 26.04
Lon [deg]: 130.75
Altitude [km]: 604.32
1.143 2.79 7.750 ×1080.26
218 November 2018
Lat [deg]: 25.11
Lon [deg]: 130.53
Altitude [km]: 604.05
1.129 0.92 2.556 ×1080.23
Aerospace 2019,6, 108 26 of 38
Figure 12. Geographic locations of the CPD missions conducted on November 2018.
3.2.2. DLP Instrument
During phase 1 of the mission (LEOP or early operations), the DLP was operated in the real-time
mode similarly to the CPD. The data were transmitted to the ground station at Kyutech when
the satellite was in view and after it had received the command for the DLP real-time operation.
During these tests, the three DLP resolution modes were selected. However, in all cases, most of the
received data packets contained zeros. Certain data packets containing data from the DLP instrument
were received among the zero values packets. However, the received samples were not sufficient
to reconstruct a single I–V curve. A problem with the DLP power line, which would prevent the
instrument for switching on for the required duration was suspected. This is presented in more detail
in the discussion section.
3.2.3. Magnetometer
During early operations, the magnetometer was also operated in real-time mode. The obtained
data shown in Figure 13 demonstrate that even though the instrument has been mounted on the
body of the spacecraft, the readings can be obtained with no significant noise coming from the other
spacecraft subsystems. In Figure 13, the sinusoidal oscillations are a consequence of the spinning of
the satellite. For attitude determination purposes, the magnetometer information is processed together
with the Sun sensor and gyroscope data. As for the GEO field readings, the local field at the spacecraft
position is obtained by using the calibrated model of the instrument. This calibration process and
the comparison of the results with geomagnetic field models like the World Magnetic Model [
have to be discussed in a separate publication in order to be able to explain it in sufficient detail to
be reproducible. Further readings of the magnetometer were not always possible (the instrument
returned only zeros), which is discussed below.
3.2.4. Material Mission (MM)
Data from the material mission have shown that it is possible to obtain information regarding
the CTE change of each of the samples from strain and temperature sensors. A comparison with the
preliminary ground tests and data measured in orbit for Sample 3 (described in Section 2.3) is shown
in Figure 14. The data show that CTE does not remain constant and is dependent on the temperature
changes as well as the carbon fiber direction.
Aerospace 2019,6, 108 27 of 38
Figure 13.
Magnetometer measurements obtained in the instrument frame of reference (
) and in the
Ten-Koh frame of reference (
) for a period of 220 s. The discontinuities come from loss of data
packets in the downlinking process during a communication session between the satellite and Kyutech
ground station.
Strain was measured in two orthogonal, in-plane directions in each sample. While manufacturing
the samples with embedded strain gauges, the alignment of the gauges with respect to the direction
of the carbon fibers was also shown to modify the readings of the applied strain. The preliminary
results from the MM appear to match the previous tests [
]. To have a complete understanding of the
material degradation, a correlation of the orbital data with the ground calibration information will be
performed. This will be included in a future publication regarding the material mission results.
Figure 14. Material mission data obtained between December 2018 and January 2019. Estimated CTE
from orbital data compared versus the ground test data for Sample 3.
4. Discussion
4.1. On-Orbit Operations History
After the launch, data started to be acquired according to the planned mission phases. The first
action was the confirmation of the beacon signal reception in several parts of the world. After that,
the operations from the Kyutech MGS started by testing all of the subsystems individually (LEOP
phase) during the three weeks following the launch. In this period, the main payloads were turned on
for a short time to verify the power consumption and that the different experimental data can be read
from the different sensors in real-time mode.
During the LEOP phase, operation training and planning were conducted with the Ten-Koh
operators team. For three weeks, the students involved in the operations developed procedures and
scheduling tasks that were tested and verified first with the EM model of the satellite before being
executed in orbit. This procedure of testing with the EM model allowed the correction of some ground
station errors, before commanding the satellite. For example, an error in the pointing of the antenna
was corrected to enable more accurate satellite tracking and improve the downlink data throughput.
Aerospace 2019,6, 108 28 of 38
Another corrected error refers to software bugs in the commanding interface application, which were
solved by an upgrade of the software interface.
Once all ground systems were tested and ready, and after the confirmation of the operation status
of the different Ten-Koh subsystems, the nominal operations phase started on 19 November 2018.
The first main mission was to operate the CPD instrument for a total duration of 4 min 43 s.
CPD and ARASE Coordinated Observations
Because the CPD instrument is the largest electrical load on-board, using power with 8.2 W,
its operating time was limited to at most 5 min twice every orbit. This high power consumption is
exacerbated by the fact that the region of operation of the CPD is mainly located at high latitudes
and around the polar caps, where the satellite can enter the eclipse region while operating the CPD.
For the specific case of the first CPD mission, on 19 November 2018, the mission was executed over
the north pole starting from a latitude of 82.2 degrees and a longitude of
165.28 degrees to a final
latitude of 70.98 degrees and a longitude of 126.63 degrees, a zone that was inside Earth’s shadow at
that time. During this first mission, data from the power consumption and battery usage (discharge)
were obtained, which allowed for the planning of longer duration CPD missions later.
In total, CPD observations were performed between 19 November 2018 and 19 March 2019,
generating approximately 5.9 MB of data. The datasets from the first 12 missions (November to
January) have been retrieved completely from the ECU SD memory, which is equivalent to 0.68 MB or
11.5% of the 5.9 MB. Because the battery and the power line that feeds the CPD have been performing
well, the duration of the CPD missions was extended, with the longest one lasting a total of 100.5 min.
This was equivalent to 201 individual readings from the CPD-CMOS and Liulin instruments. Table 3
shows that one of the challenges in the satellite development is the volume of data generated by the
main payload instruments. The storage of such amount of data was addressed by selecting a storage
device with 2 GB capacity. However, the communication bandwidth limitation remains the bottleneck
when retrieving experiment data.
Based on the characteristics of the Ten-Koh orbit and the CPD detectors energy range,
a collaboration with the ARASE (ERG) satellite [
] started from December 2018. Conjunction
operations for the observation of high-energy electrons and ions were deducted feasible in locations
when the difference between latitude and longitude of both satellites footprint at 100 km altitude is
within five degrees. An example of the conjunction points between both satellites is presented in
Figure 15. The ARASE team provided the conjunction points of both satellites by the use of the ARASE
orbit prediction tool and the latest TLE available for Ten-Koh. Then, the Ten-Koh team produced the
corresponding set of commands and schedule the time of operations for commanding the satellite.
The satellite performed the joint observations with ARASE satellite when the on-board time matched
the epoch specified in the uplinked command. The results from the cooperation with the ARASE
satellite will be presented in a subsequent publication because they do not fall in the scope of this paper.
Aerospace 2019,6, 108 29 of 38
Figure 15.
Ten-Koh and ERG (ARASE) conjunction points: (
) world map of the conjunction points for
the period between 4 to 30 December 2019, showing both satellites footprints; and (
) example of two
conjunction points computed for the period between 21 February 2019 and 22 February 2019.
4.2. Mission Operations Issues and Status
4.2.1. CPD Operation
The CPD was tested successfully during the LEOP phase of the satellite operations. The five
datasets obtained from the Liulin detector over Japan were measured with an average exposition
time of 15.2 s. In comparison, the exposition time for the normal mode missions was 30 s in average.
The reduced exposure time caused the early operation (real-time) data to receive fewer counts, and
less total dose rate and flux than the normal mode missions data. Variation in the dose and flux
for the real-time data comes from two data series (12 November at 05:07:29 UTC and 18 November
at 04:34:08 UTC) with most of the hits in the low deposited-energy channels of the detector. The
main reason for the decrease in dose rate and flux in the mentioned datasets relates to the rotation
of the satellite and the misalignment of the magnetic field lines with respect to the normal direction
of the Liulin detector. The effect of the satellite rotation was perceived from the beginning of the
operations during the communication sessions with the Kyutech MGS. The rotation of the satellite
makes the beacon signal vary in intensity during communication sessions, which was observed during
all communication session operations. This highlights the fact that schedule and resource constraints
of the project have had a measurable effect on the scientific return of the mission. A similar pattern is
Aerospace 2019,6, 108 30 of 38
expected for other satellites that are launched as secondary (piggyback) payloads, without the ability
to choose the launch date.
The Liulin results show an increase in dose rate and flux in regions where the ambient models
predict the presence of electrons with more than 2.2 MeV and protons with more than 30 MeV.
These energy values correspond to the energy required to penetrate the shielding of Ten-Koh external
structure for each particle type. Had the instrument been placed with an unobstructed view of
deep-space, lower energies could have also been registered. This, however, would have necessitated a
complete redesign of the satellite structure that was not feasible from schedule point of view.
4.2.2. DLP Failure
A problem with the power line that feeds the DLP measurement system was encountered,
making it impossible for the 5 V line to sustain its nominal output level. The root cause of the problem
was identified as activation of the overcurrent protection due to the high-frequency transient from the
inrush current of the DLP DC/DC converter. During the thermal-vacuum (TVAC) tests, the problem
appeared for the first time in the EM model of the satellite. The issue was solved by reducing the
amplitude of the inrush current through the use of an in-series resistor with the power line output
from the EPS. However, once in orbit, it is suspected that changes in resistance of the used resistor
(less than 1
resistor value) triggered by temperature variations with respect to ambient conditions
on-ground made the problem reappear.
Comparison tests performed with the EM and FM spare models of the DLP and the EM regulation
board from the EPS showed a discrepancy in the inrush current of the DLP DC/DC converter between
the EM and FM versions. Even though the difference in the inrush current from EM to FM translates
to a difference of the in-series resistor of 0.3
, it is high enough to trigger the OCP in the FM EPS
DLP power line. During the calibration tests and the development of the flight software, the DLP
instrument was powered by a workbench power supply adjusted to the same current limit as the
DLP line from the EPS (1.5 A). The nominal current consumption was 0.27 A for both the EM and FM
hardware. This caused the inrush current issue not to be identified at an earlier stage.
If re-work of the satellite were possible, a power regulation board (REG_PL) that accounts for the
non-linear nature of different loads should be designed to avoid similar issues. As for addressing the
issue of the unit currently in orbit, temperature variations will cause the in-series resistor to change its
value. Therefore, it is possible that the DLP instrument could be operated at certain times during the
lifetime of the spacecraft, as the orbit will naturally evolve and the thermal environment inside the
satellite will vary accordingly.
4.2.3. SD Card and Magnetometer Readings Failure
The failures in reading the SD card and the magnetometer were caused by the controller board,
which includes a single PIC microcontroller. The same PIC is used to read the telemetries of the 12
solar panels (voltage, current and temperature) that are also used for attitude determination purposes.
This solution has been followed in order to synchronize the readings of all the ADS telemetries.
However, managing so many devices with a single PIC strained the available hardware resources
of the microcontroller. The PIC did not have in-built hardware support for all the data buses that it
was required to operate simultaneously. Thus, I2C communication with the OBC was implemented
making use of the in-built hardware in the microcontroller, while SPI communication with the sensors
and the SD card was implemented in software.
Solar panel and magnetometer board ADCs are sequentially polled to keep their readings
synchronized. The acquired data are either stored on the SD card or directly transferred to the
OBC for near real-time downlink. When implementing the SPI communication protocol entirely in
software, the timing of the clock and the data signals needs to be maintained in order to keep the
system operational. Changes in the timing of these signals turned out to be critical when operating the
sensor polling sequence, and have ultimately caused the entire SPI bus to suffer partial or complete
Aerospace 2019,6, 108 31 of 38
outages for prolonged periods of time. The changes in this timing could have been caused by either
radiation effects, or by changes in temperature of the PIC and its local oscillator. Even though all of the
ADS devices have been connected to a single SPI data bus, a propagation of a hardware failure of one
of these devices can be ruled out because the devices have been isolated from each other.
The SD card failed after the first test during the LEOP phase with no possibility of recovering it
later. The magnetometer could be read only once during the LEOP phase, and again after 4 March 2019.
Between 4 and 17 March 2019, the magnetometer readings operated nominally in the real-time
mode that did not make use of the SD card. However, operation in real-time mode means that the
magnetometer could only be operated when the satellite was passing over the Kyutech ground station.
Only a power cycle of the system could restore the ADS operation, which was confirmed during
ground tests. However, because the EPS and the ADS controller boards have not been designed to be
powered off once the satellite began operating, this solution cannot be implemented in-orbit. All other
subsystems can be power cycled, including all the payloads, both COMMS boards, and the OBC
board. Even though the EEPROM in the OBC PIC has encountered errors during operation, the system
has resumed nominal operations after a power cycle. In the case of the payloads, the ECU and MM
controllers are kept off most of the time, and only receive power when an experiment is scheduled
for execution. No errors have been encountered in these microcontrollers since the launch. Based on
this microcontroller operation evidence, the lack of power cycle capability in the EPS and ADS boards
made the microcontrollers of these subsystems more susceptible to failure, and limited the number of
means to recover them.
4.2.4. OBC EEPROM Errors
Errors in the microcontroller EEPROM in the OBC were detected during the operation. The errors
made changes in the log register of the commands processed by the OBC and their origin is suspected
to be related to single event upsets (SEU). The OBC operation was always recovered by applying a
power cycle from the EPS and then updating the on-board time of the satellite by a ground command.
During the period of the last days of December 2018 and the first days of January 2019, more resets in
the OBC were carried out than in the preceding weeks. However, none of these resets occurred during
the execution of one of the main missions. Most of the errors and OBC resets in this period caused at
most a change in the satellite time and the loss of the last received commands.
Besides the OBC EEPROM errors, similar errors were observed in other microcontrollers’
EEPROMs. In particular, the EPS microcontroller experienced a change in the SD card memory
address value stored in the EEPROM. This value is used by the EPS microcontroller to know the last
used location of the SD card, where the most-recent housekeeping data have been stored. The solution
for the EPS EEPROM issue was a regular update of the SD address for saving of the data following a
ground command.
Commercial CMOS-EEPROMs have been shown to be susceptible to SEU while they are powered
on [
]. Even though the HEF-EEPROMs (high-endurance flash) on-board Ten-Koh use a different
technology, it is suspected that they might be sensitive to SEU in the same way. Because the
microcontroller EEPROM is used to store commands and/or update parameters during the execution
of the main software routine, changes in bits in the EEPROM could potentially alter and even stop the
correct operation of the microcontroller completely. A detailed investigation of the radiation effects
on the HEF-EEPROM of the PIC16F877 is being conducted and the results will be published in a
subsequent publication.
4.2.5. Spacecraft Status
The Ten-Koh spacecraft was continuously operated daily since its launch (29 October 2018) until
19 March 2019, when a loss of communication between the spacecraft and the Kyutech ground station
occurred. The ground-track of the last orbits of the satellite where the failure occurred are shown in
Figure 16. The last data from its beacon were received over Greece on 18 March 2019 between 23:12:33
Aerospace 2019,6, 108 32 of 38
and 23:25:43 UTC. The decoded data indicated that all the main housekeeping parameters were inside
their operating ranges, as shown in Figure 17. According to the same beacon data, the satellite was
in the expected mission mode, i.e., nominal CPD mission scheduled between 22:00 on 18 March and
00:20 on 19 March UTC. The mission was supposed to finish when the satellite was coming from
above the South Pole and heading northwest towards the southeast of South Africa. After finishing
the CPD mission, Ten-Koh traveled for three more orbits, including two of them over the SA anomaly,
before the pass over Kyutech ground station on 19 March 2019 at 13:31 JST (19 March 2019 at 04:31
UTC) when no signal was received. Irrespective of the anomaly that made the satellite fail, it occurred
during the two orbits following the CPD mission and before the pass over the Kyutech ground station.
A review of the geomagnetic activity parameters from the NOAA Space Weather Prediction Center
between 15 March 2019 and 18 March 2019 revealed that the Kp-index, an average of K-indices from a
network of geomagnetic observatories, showed above the threshold value of 4 on 17 March 2019. As
shown in Figure 18, Kp = 4 was observed between 21:00 on 16 March 2019 and 00:00 on 17 March 2019
UTC, and increased to Kp = 5 between 00:00 and 03:00 on 17 March 2019 UTC. Kp then decreased
to Kp = 4 between 03:00 and 09:00 on 17 March 2019 UTC. The Kp-index with values bigger than 4
indicates a geomagnetic storm activity.
The Ten-Koh anomaly occurred approximately 50 h after the geomagnetic activity (between 18:00
on 18 March 2019 UTC and 03:00 on 19 March 2019 UTC). This shows a direct correlation between the
anomaly and the presence of a geomagnetic storm. Space weather can produce temporal variations in
the trapped radiation [
], which in turn can cause SEE and charging (surface and internal) that result
in anomalies experienced by satellites.
The main hypothesis for the failure of the satellite was an error triggered by a single event effect
(SEE), which was previously observed as errors in the in-built EEPROM of several microcontrollers.
Note that such events were expected to take place over South Atlantic Anomaly due to presence of
high-energy protons (in excess of 30 MeV) that could penetrate the spacecraft structure.
As in the case of the readings from the magnetometer that were stopped and then recovered
after some time, the satellite began to transmit its beacon again after a solar storm that occurred
on 10 May 2019 and that arrived at Earth between 14 and 16 May 2019. This correlation between
and increased number of particles in LEO caused by the solar storm, and recovery of the satellite
functionality corroborates the hypothesis that it was originally disabled by radiation damage. It is
thought that the bit changes in the in-built EEPROM or microcontroller were undone by subsequent
single event upsets so that the satellite could resume operations. It is also possible that the satellite was
originally disabled by a single event latch-up, which did not increase the OBC power consumption
enough to trigger the over-current protection. Subsequent latch-ups could have eventually caused the
OCP trip, and the OBC to reboot and resume operations. This, however, is not the most likely scenario
because multiple reboots were attempted from the ground but to no avail.
Aerospace 2019,6, 108 33 of 38
Figure 16.
Ten-Koh orbit ground-tracks when the satellite is suspected to have failed. Following a
beacon reception in Greece, Ten-Koh flew over the SA region in two consecutive orbits (shown in
orange) before passing over the Kyutech ground station (yellow mark) in Japan on 19 March 2018.
Figure 17.
Last decoded beacon received from Ten-Koh over Greece on 19 March 2019. Note that “Date”
and “Time” correspond to the time when the beacon was decoded, not when it was transmitted.
Aerospace 2019,6, 108 34 of 38
Figure 18.
Estimated planetary index during the time the Ten-Koh anomaly occurred on 19 March 2019
(showed with the red bars).
At the time of writing, the satellite regularly transmits its beacon that includes correct on-board
time and temperature data, and receives commands. However, conducting further experiments is
currently impossible because the microcontroller governing the EPS is not reacting to commands or
providing any telemetry. Because of the issues that the EPS controller has encountered previously, it is
believed that the PIC’s EEPROM is corrupted, which is preventing the reception of commands from the
OBC via serial communication buses. The microcontroller reacts to hardware interrupts that are used to
trigger a satellite reset, which confirms that the PIC itself is still functioning. Currently, the operations
continue to try and recover full functionality of the spacecraft, which is plausible given the history of
the previous failures.
5. Conclusions
The Ten-Koh satellite, launched on 29 October 2018, was the second satellite developed in a series
of two, similarly-sized, low-cost and rapidly-developed small satellites from the Okuyama laboratory
of the Kyushu Institute of Technology. The satellite was used to collect data pertaining to degradation
of materials in space and high-energy particle fluxes in low-Earth orbit during the solar minimum.
Preliminary analysis of the data has shown that the main missions were able to acquire data within
the expected ranges, except for the DLP that could not be operated due to the described power line
failure. The CPD instrument has detected electrons, protons and galactic cosmic rays, which provides
information about the radiation environment inside the spacecraft. In the case of the magnetometer,
it has been shown that the selected design, where the instrument is attached directly on the top of the
spacecraft without a boom, is able to produce measurements of the geofield that serve for attitude and
science purposes. Above all, the in-orbit data have shown that the university-built satellite worked
as intended, which is a success in its own right due to the limited resources that were available for
its development.
The joint observations with the ARASE satellite show that the high-energy electrons mission
on-board Ten-Koh has contributed to the field of particle science. Moreover, this mission has shown
that small-satellites can be used as a complement to other missions investigating space environment.
Aerospace 2019,6, 108 35 of 38
A variety of failures, relating to both the satellite subsystems and the ground segment have been
described. The most critical ones have been associated with triggering of over-current protection,
degradation of EEPROM memory due to space radiation, and single-event upsets. These should be
taken into account when developing future low-cost space missions, which will face similar challenges
to those described in this paper. The best means of recovering from such failures has been found to be
a power cycle of a unit, which should be implemented in future missions.
The abundance of radiation-related failures, with consequences as severe as disabling the entire
satellite, indicates that understanding of the space environment and its effects remains vital. This shows
that the mission objectives of Ten-Koh served a valid scientific purpose, and should be pursued by
more missions. Space environment effects are especially important in the era where small satellites are
proliferating because such spacecraft do not use radiation-hardened parts and thus are likely to suffer
similar issues as the satellite described herein. Thus, the lessons learned described in this paper should
be taken into account when developing similar satellites.
Lastly, the interplay between programmatic constraints and scientific return has been highlighted.
One notable example of this is how discarding attitude control in order to meet the resources and
schedule constraints has caused satellite rotation to be visible in CPD readings. Trading off cost,
schedule, risk as well as return of the mission will be particularly important for small satellites,
which cannot afford the luxury of postponing launches in order to meet their original scientific
objectives. This is a large shortcoming of relying on piggyback launch opportunities, and space
agencies worldwide should consider offering more flexible launch slots in order to maximize the return
they get from small satellite missions. If the launch date cannot be postponed, spacecraft developers
ought to consider reducing the scope of their mission by discarding certain objectives in order to satisfy
the schedule constraint.
Author Contributions:
Conceptualization, I.F., A.A.L., J.G.-L., R.R., R.M. and K.-i.O.; methodology, I.F. A.A.L.,
J.G.-L., K.-i.O., P.S., D.H. and T.D.; validation, I.F., A.A.L., S.A.B., J.G.-L, R.R., R.M., K.-i.O., D.F., M.M. (Misuzu
Matsuoka), N.U., R.K., M.M. (Masayuki Miyazaki), N.Y., K.H., F.A., J.J.R., M.E.K., K.C., T.U., P.S., D.H., T.D.,
S.T. and R.D.; formal analysis, I.F., A.A.L., J.G.-L., K.-i.O., P.S., D.H. and T.D.; investigation, I.F., A.A.L. and
K.-i.O.; resources, K.-i.O., P.S. and D.H.; data curation, I.F., A.A.L., K.-i.O., P.S., D.H. and T.D.; writing—original
draft preparation, I.F. and A.A.L.; writing—review and editing, I.F., A.A.L., J.G.-L., R.M., K.-i.O., S.T. and R.D.;
visualization, I.F, A.A.L., R.M.; supervision, K.-i.O.; project administration, K.-i.O.; funding acquisition, K.-i.O.,
P.S. and D.H.; and software, I.F., A.A.L., P.S. and D.H.
The authors would also like to extend thanks and appreciation to the Oita Prefectural government,
the Oita Prefectural Organization for the Industry Creation, and the Working group for Ten-Koh development
“Oita Challenger", for their financial and technical support in all the process of the satellite mission project.
First and foremost, the authors would like to extend their gratitude to JAXA for offering the
launch opportunity to the Ten-Koh team. Isai Fajardo appreciates the support for the scholarship he receives
from the Ministry of Education, Culture, Sports, Science and Technology of Japan (MEXT). Aleksander A. Lidtke
would like to acknowledge the funding he received from the Ministry of Education, Culture, Sports, Science and
Technology of Japan. Jesus Gonzalez-Llorente would like to thank the support for the scholarship he receives
from the Ministry of Education, Culture, Sports, Science and Technology of Japan (MEXT). Rafael Rodriguez also
appreciates the support for the scholarship he receives from the Ministry of Education, Culture, Sports, Science
and Technology of Japan (MEXT). Rigoberto Morales extends his gratitude to the Mexican Council of Science and
Technology “CONACYT” for the fully funded scholarship he received.
Conflicts of Interest:
This is an original submission that has not been published before and that is not currently
under review consideration for publication elsewhere.The authors declare no conflict of interest. The funders
had no role in the design of the study; in the collection, analyses, or interpretation of data; in the writing of the
manuscript, or in the decision to publish the results.
Aerospace 2019,6, 108 36 of 38
The following abbreviations are used in this manuscript:
ADC Analog to Digital Converter
ADS Attitude Determination Subsystem
COMM Telecommunications subsystem of Ten-Koh
CCU Communications Control Unit (of Ten-Koh COMM)
CFRP Carbon Fiber Reinforced Polymer
CFRTP Carbon Fiber-Reinforced Thermoplastic
ECU Experiment Control Unit
EPS Electrical Power Subsystem
GFRP Glass Fiber Reinforced Polymer
GNSS Global Navigation Satellite System
GOSAT Greenhouse gases Observing SATellite
GPIO General-Purpose Input Output
CPD Charged Particle Detector, one of Ten-Koh payloads
COTS Commercial off-the-shelf
DLP Double Langmuir Probe, one of Ten-Koh payloads
I2C Inter-Integrated Circuit (protocol)
JAXA Japan Aerospace Exploration Agency
Kyutech Kyushu Institute of Technology
LEO Low Earth Orbit
LEOP Launch and Early phase (of the Ten-Koh mission)
MCU Main Controller Unit (of Ten-Koh OBC)
MGS Main Ground Station
OBC Onboard computer
OCP Over-Current Protection
PCB Printed Circuit Board
PIC Programmable Integrated Circuit
PL Payload
RTC Real Time Clock
SPI Serial peripheral interface (protocol)
TTL Transistor–transistor logic
UHF Ultra-High Frequency
UCP Ultracapacitor
UV Ultraviolet
Sihver, L.; Kodaira, S.; Ambrožová, I.; Uchihori, Y.; Shurshakov, V. Radiation Environment Onboard
Spacecraft at LEO and in Deep Space. In Proceedings of the 2016 IEEE Aerospace Conference,
Big Sky, MT, USA, 5–12 March 2016; pp. 1–9. [CrossRef]
Suparta, W.; Gusrizal. The Variability of Space Radiation Hazards towards LEO Spacecraft. In Journal of
Physics: Conference Series 539 (October 2014): 012023; IOP Publishing: Bristol, UK, 2014. [CrossRef]
Chen, L.; Thorne, R.M.; Li, W.; Bortnik, J.; Turner, D.; Angelopoulos, V. Modulation of Plasmaspheric Hiss
Intensity by Thermal Plasma Density Structure. Geophys. Res. Lett. 2012,14,39. [CrossRef]
Cucinotta, F.A.; Cacao, E.; Kim, M.H.Y.; Saganti, P.B. Cancer and Circulatory Disease Risks for a Human
Mission to Mars: Private Mission Considerations. Acta Astronaut. 2018,13. [CrossRef]
Miyake, S.; Kataoka, R.; Sato, T. Cosmic Ray Modulation and Radiation Dose of Aircrews during the Solar
Cycle 24/25. Space Weather 2017,15, 589–605. [CrossRef]
Kirby, K.; Bushman, S.; Butler, M.; Conde, R.; Fretz, K.; Herrmann, C.; Hill, A.; Maurer, R.; Nichols, R.; Ottman,
G. Radiation Belt Storm Probe Spacecraft and Impact of Environment on Spacecraft Design. In Proceedings
of the 2012 IEEE Aerospace Conference, Big Sky, MT, USA, 3–10 March 2012; pp. 1–20. [CrossRef]
Stratton, J.M.; Harvey, R.J.; Heyler, G.A. Mission Overview for the Radiation Belt Storm Probes Mission.
Space Sci. Rev. 2013,179, 29–57. [CrossRef]
Aerospace 2019,6, 108 37 of 38
8. Angelopoulos, V. The THEMIS Mission. Space Sci. Rev. 2008,141, 5–34. [CrossRef]
Sharma, A.S.; Curtis, S.A. Magnetospheric Multiscale Mission. In Nonequilibrium Phenomena in Plasmas;
Astrophysics and Space Science Library; Burton, W.B., Kuijpers, J.M.E., Van Den Heuvel, E.P.J.,
Van Der Laan, H., Appenzeller, I., Bahcall, J.N., Bertola, F., Cassinelli, J.P., Cesarsky, C.J., Engvold, O.,
et al., Eds.; Springer: Dordrecht, The Netherlands, 2005; pp. 179–195.[CrossRef]
Nakamura, Y.; Fukuda, S.; Shibano, Y.; Ogawa, H.; Sakai, S.I.; Shimizu, S.; Soken, E.; Miyazawa, Y.; Toyota, H.;
Kukita, A.; et al. Exploration of Energization and Radiation in Geospace (ERG): Challenges, Development,
and Operation of Satellite Systems. Earth Planets Space 2018,70, 102. [CrossRef]
Santandrea, S.; Gantois, K.; Strauch, K.; Teston, F.; Tilmans, E.; Baijot, C.; Gerrits, D.; Team, P.I.; De Groof, A.;
Schwehm, G.; et al. PROBA2: Mission and Spacecraft Overview. Sol. Phys. 2013,286, 5–19. [CrossRef]
Nils, O.; Rune, F. Exploring Geospace from Space: The Swarm Satellite Constellation Mission. Space Res. Today
2018,203, 61–71. [CrossRef]
Xapsos, M.A.; O’Neill, P.M.; O’Brien, T.P. Near-Earth Space Radiation Models. IEEE Trans. Nucl. Sci.
60, 1691–1705. [CrossRef]
Dudziak, R.P.; Tuttle, S.L.; Okuyama, K.; Lidtke, A.A.; Fajardo Tapia, I.; Gonzales, J.D.
Practical Considerations of Integrating a Passive Thermal Control System onto Small-Satellites-the Ten-Koh
Case Study. In Proceedings of the 49th International Conference on Environmental Systems, Boston, MA,
USA, 7–11 July 2019.
Bendoukha, S.A.; Okuyama, K.I.; Bianca, S.; Nishio, M. Control System Design of an Ultra-Small Deep Space
Probe. Energy Proced. 2016,100, 537–550. [CrossRef]
Eickhoff, J. A Combined Data and Power Management Infrastructure: For Small Satellites; Springer Aerospace
Technology; Springer-Verlag: Heidelberg, Germany, 2013; pp. 51–52. ISBN 978-3-642-35556-1.
Reyes, R.; Faizullin, D.; Fajardo, I.; Hiraki, K.; Okuyama K. Nano-satellite TEN-KOH Attitude Determination
Subsystem: Architecture and Initial In-orbit Results. In Proceedings of the 32nd International Symposium
on Space Technology and Science & 9th Nano-Satellite Symposium, Fukui, Japan, 15–21 June 2019.
Gonzalez-Llorente, J.; Lidtke, A.A.; Hatanaka, K.; Kawauchi, R.; Okuyama, K.I. Solar Module Integrated
Converters as Power Generator in Small Spacecrafts: Design and Verification Approach. Aerospace
2019,6, 61. [CrossRef]
Dachev, T.P.; Bankov, N.G.; Tomov, B.T.; Matviichuk, Y.N.; Dimitrov, P.G.; Häder, D.-P.; Horneck, G. Overview
of the ISS Radiation Environment Observed during the ESA EXPOSE-R2 Mission in 2014–2016. Space Weather
2017,15, 1475–89. [CrossRef]
Dorman, L.I.N.; Iucci, A.V.; Belov, A.E.; Levitin, E.A.; Eroshenko, N.G.; Ptitsyna, G.; Villoresi, G.V.;
Chizhenkov, L.I.; Gromova, M.; Parisi, M.I.; et al. Space Weather and Space Anomalies. Ann. Geophys.
2005,23, 3009–3018. [CrossRef]
Takuji, E. A Miniature GNSS Smart Antenna for Space Applications. Available online: https://www3. (accessed on
26 August 2019).
Gosat-2 Launch Information.Pdf. Available online:
h2af40_j.pdf (accessed on 31 March 2019).
23. Bilitza, D. IRI the International Standard for the Ionosphere. Adv. Radio Sci. 2018,16, 1–11. [CrossRef]
International Reference Ionosphere-IRI. 2016. Available online:
models/iri2016_vitmo.php (accessed on 31 March 2019).
Scott, F.N. Future Trends in Satellite Systems and Their Impact on Material Choice. Mater. Today
,2, 7–10.
Ken Chang, C.; Kamaratos, E. Theoretical Studies of Radiation Effects in Composite Materials for Space Use,
NASA Contractor Report 3618. 1982. Available online:
gov/19820026459.pdf (accessed on 28 March 2019).
Dachev, T.P.; Semkova, J.V.; Tomov, B.T.; Matviichuk, Y.N.; Dimitrov, P.G.; Koleva, R.T.; Bankov, N.G.;
Shurshakov, V.A.; Benghin, V.V.; Yarmanova, E.N.; et al. Overview of the Liulin Type Instruments for Space
Radiation Measurement and Their Scientific Results. Life Sci. Space Res.
,4, 92–114. [CrossRef] [PubMed]
Chen, F.F. Langmuir Probe Diagnostics, Mini-Course on Plasma Diagnostics. In Proceedings of the
IEEE-ICOPS Meeting, Jeju, Korea, 5 June 2003; p. 42.
29. McIlwain, C.E. Magnetic Coordinates. Space Sci. Rev. 1966,5, 585–98. [CrossRef]
Aerospace 2019,6, 108 38 of 38
Chulliat, A.; Macmillan, S.; Alken, P.; Beggan, C.; Nair, M.; Hamilton, B.; Woods, A.; Ridley, V.; Maus, S.;
Thomson, A. The US/UK World Magnetic Model for 2015–2020: Technical Report; National Geophysical Data
Center, NOAA: Silver Spring, MD, USA, 2015. [CrossRef]
Dong, C.; Li, K.; Jiang, Y.; Arola, D.; Zhang, D. Evaluation of Thermal Expansion Coefficient of Carbon
Fiber Reinforced Composites Using Electronic Speckle Interferometry. Opt. Express
,26, 531. [CrossRef]
Arase (ERG) Geospace Probe|Spacecraft. ISAS. Available online:
spacecraft/current/erg.html (accessed on 21 April 2019).
Harboe Sorensen, R.; Muller, R. Radiation Testing of UV EPROMs, Flash EPROMs and EEPROMs for
Space Applications. ESA Report ESA-QCA0076TS. 1996. Available online:
webDocumentFile?id=1026 (accessed on 3 August 2019).
Baker, D.N.; Erickson, P.J.; Fennell, J.F.; Foster, J.C.; Jaynes, A.N.; Verronen, P.T. Space Weather Effects in the
Earth’s Radiation Belts. Space Sci. Rev. 2017,214, 17. [CrossRef]
2019 by the authors. Licensee MDPI, Basel, Switzerland. This article is an open access
article distributed under the terms and conditions of the Creative Commons Attribution
(CC BY) license (
... For example, the FORMOSAT-5 [37] has three sensors (LP, RPA, IDM) that are run sequentially at different sampling rates to measure ionospheric plasma in a cycle mode. Also, the GOSAT-2 (greenhouse gases observing satellite-2) mission [38] has a variety of plasma sensors such as a charged particles detector, a double LP, and a magnetometer, all of them running sequentially and storing the acquired data in secure digital (SD) cards-the oldest kind of memory card, with up to 2GB memory capacity. Given the limitations for servicing satellite hardware in orbit, it would be desirable to gather redundant plasma data using a set of similar sensors or different sensors that generate data that to some extent overlap. ...
... quasistatic bias voltages. The minimum average speed of a satellite with a stable orbit in LEO is about 7.8 km s −1 [21,38] which corresponds to an ion energy E o = 0.32N i eV, where N i is the mass number of the ionized species. Therefore, each species will show in the RPA data as a peak with a different average ion energy. ...
Full-text available
This review details the state of the art in in-space plasma diagnostics for characterizing the Earth’s ionosphere. The review provides a historical perspective, focusing on the last 20 years and on eight of the most commonly used plasma sensors—most of them for in situ probing, many of them with completed/in-progress space missions: (a) Langmuir probes, (b) retarding potential analysers, (c) ion drift meters, (d) Faraday cups, (e) integrated miniaturized electrostatic analysers, (f) multipole resonance probes, (g) Fourier transform infrared spectrometers, and (h) ultraviolet absorption spectrometers. For each sensor, the review covers (a) a succinct description of its principle of operation, (b) highlights of the reported hardware flown/planned to fly in a satellite or that could be put in a CubeSat given that is miniaturized, and (c) a brief description of the space missions that have utilized such sensor and their findings. Finally, the review suggests tentative directions for future research.
... Low-earth orbit (LEO) satellites [1][2][3][4], revolving at an altitude between 160 to 2,000 km, can offer faster communications with low latency and higher bandwidth at low cost. In addition, a constellation of LEO satellites-Starlink, LeoSat, Boeing, CYGNSS, and Telesat LEO which provide various satellite services such as remote sensing, weather monitoring, internet of things (IoT), and satellite navigation [5,6]-can attain continuous and global coverage for various communication systems. ...
... An effective refractive index neff in the troposphere and stratosphere can be calculated from the complex refractivity N via (2) where N`and N" are real and imaginary values of N [17]. In the ionosphere (i.e., cold/magnetized plasmas), the effective refractive index can be calculated by using Appleton-Hartree equation. ...
Full-text available
We propose a novel method to calculate the electromagnetic (EM) wave propagation from low-earth orbit satellite (LEO) to a ground station based on the physical optics (PO), ray tracing technique, and geometric optics (GO) considering interpolated atmospheric environments. Our method includes the reflector antenna analysis using PO, the interpolation of the meteorological data using PCHIP and Kriging interpolation, transmission analysis using ray tracing and geometrical optics. Tropospheric and stratospheric environments are modeled using meteorological data–air pressure and temperature, relative humidity, and rain rate–measured at 9 different radiosonde observatories in and around South Korea. Furthermore, we utilize Piecewise Cubic Hermite Interpolating Polynomial (PCHIP) and Kriging-exponential methods for vertical and horizontal interpolations of the raw meteorological data, respectively. Hence, the interpolated atmospheric environments are amenable to the best use of ray tracing technique and GO. Subsequently, effective refractive indices of the stratified media can be extracted via millimeter-propagation-model93. The simplified Appleton-Hartree equation characterizes the ionospheric environment. Considering a sunsynchronous orbit satellite passing through South Korea, we calculate atmospheric attenuation, boresight error, received power, and compensation angle of satellite antenna for various conditions.
... Tin seems to be more likely to grow over time whiskers in a vacuum, providing a short circuit path between metal-plated surfaces [147]. However, many dedicated SmallSats of this category are offering an incredible job to explore space environments, which will enable building more robust and reliable systems [148]. In addition, at the time of left-off vibration is the sudden application of 9.2 million pounds of thrust to the satellite [146]. ...
Full-text available
The Small Satellite (SmallSat) industry has recorded incredible growth recently. Within this class, among Mini-, Micro-, and Nanosatellites, the Cube Satellite (CubeSat) is primed for an explosion of growth. These satellites are fascinating for remote sensing, earth observation, and scientific applications. Remarkable attention from the space operators makes it valuable because of its low cost, cubic shape, less manufacturing time, lightweight, and modular structure. Among the various subsystems comprising the SmallSat, the Electrical Power System (EPS) is the most crucial one because unreliable power supply to the rest is most of the time detrimental to the mission. The EPS is formed by electrical sources, storage units, and loads, all interconnected via different power converters, the operation of which must be closely orchestrated to accomplish efficient use of photovoltaic power, optimal battery management, and resilient power delivery. At the same time, the EPS design must address a series of challenges such as size restrictions, high power density, harsh space environments (e.g., atomic oxygen, radiation, and extreme temperatures) which significantly impact the EPS electrical and electronic equipment. In terms of power systems, a SmallSat EPS can be considered a space microgrid owing coordination and control of distributed generation (DG), storage and loads in a small-scale electrical network. From this point of view, this paper reviews and explores SmallSat microgrid’s research developments, energy transfer and architectures, converter topologies, latest technologies, main challenges, and some potential solutions which will enable building a more robust, resilient, and efficient EPS. The research gaps and future developments are underlined before the paper is concluded.
... Taking into consideration past small satellite designs launched into orbit, the right prismatic geometric shape was selected for the overall configuration of the satellite being designed, since this shape provides the best use in terms of volume with minimal volumetric waste, as described in works such as Aborehab et al. [47], compared to other geometric shapes, e.g., cylindrical, hexagonal, octagonal, and others, besides being easier to fabricate and assemble. Taking inspiration from past designs, e.g., [48], a central load bearing assembly was added to the structure to act as its backbone and to withstand the highly dynamic loads imposed on the satellite during the launch phase of its operational life, as will be seen in Section 3.4, describing the results of quasi-static launch loads analyses. Figure 1a shows the conceptual subsystem configuration after performing several component distribution designs. ...
Full-text available
Mass reduction is a primary design goal pursued in satellite structural design, since the launch cost is proportional to their total mass. The most common mass reduction method currently employed is to introduce honeycomb structures, with space qualified composite materials as facing materials, into the structural design, especially for satellites with larger masses. However, efficient implementation of these materials requires significant expertise in their design, analysis, and fabrication processes; moreover, the material procurement costs are high, therefore increasing the overall program costs. Thus, the current work proposes a low-cost alternative approach through the design and implementation of geometrically-shaped, parametrically-defined metal perforation patterns, fabricated by standard processes. These patterns included four geometric shapes (diamonds, hexagons, squares, and triangles) implemented onto several components of a structural design for a conceptual satellite, with a parametric design space defined by two scale factors and also two aspect ratio variations. The change in the structure’s fundamental natural frequency, as a result of implementing each pattern shape and parameter variation, was the selection criterion, due to its importance during the launcher selection process. The best pattern from among the four alternatives was then selected, after having validated the computational methodology through implementing experimental modal analysis on a scaled down physical model of a primary load-bearing component of the structural design. From the findings, a significant mass reduction percentage of 23.15%, utilizing the proposed perforation concept, was achieved in the final parametric design iteration relative to the baseline unperforated case while maintaining the same fundamental frequency. Dynamic loading analysis was also conducted, utilizing both the baseline unperforated and the finalized perforated designs, to check its capability to withstand realistic launch loads through applying quasi-static loads. The findings show that the final perforated design outperformed the baseline unperforated design with respect to the maximum displacements, maximum Von Mises stresses, and also the computed margin of safety. With these encouraging outcomes, the perforated design concept proved that it could provide an opportunity to develop low-cost satellite structural designs with reduced mass.
... On-board data handling subsystem (OBDH) is connected to all subsystems and manages all the processes or tasks as required. Figure 1 shows OBDH connects to the communications subsystem (COMS), electrical power subsystem (EPS), attitude determination and control subsystem (ADCS), thermal subsystem, and payload [10]. These subsystems may have sensors and actuators to give input to OBDH and allow it to make positional adjustments to achieve mission requirements. ...
Full-text available
p>CubeSat is a small-sized satellite that provides a cheaper option for the manufacturer to have a fully operational satellite. Due to its size, CubeSat can only generate limited power, and this will restrict its functionality. This research aims to improve CubeSat’s power consumption by implementing the dynamic voltage and frequency scaling (DVFS) technique to on-board and data handling subsystem (OBDH). DVFS will find the best operating frequency to execute all of OBDH’s task. This paper explains how we determined the task set, representing all routine tasks performed by OBDH during normal operation mode. We have simulated the task set using two DVFS algorithms, static earliest deadline first (EDF) and cycle conserving edf (CC EDF). The result shows that both scheduling algorithms give a similar result to our task set. However, when the scheduler is configured as non-preemptive, the simulator failed to schedule the critical task. It means that the system fails to work as intended. Therefore, we conclude that we need to implement mixed-criticality scheduling to prevent critical tasks from being aborted by the system.</p
... Кюшу, Япония от проф. К. Окуяма и неговия екип [12,13]. Той беше успешно изстрелян на 29 октомври 2018 г. на височина 623 km и наклон на орбитата от 98°. ...
... In the literature, there are some studies which discuss the space weather effects on the satellite components [6]. In a study investigating the ionizing dose effects of Globalstar M070 manufactured by Globalstar Inc. from Covington, Louisiana, U.S., Razaksat of Astronautic Technology Sdn Bhd from Malaysia and MKA-FKI 1 satellites produced by Kotelnikov Radio Technology and Electronics Institute from Russia on aluminum protection, it was observed that, as the thickness of the aluminum shield increases, the effect of electrons decreases, and the effect of protons does not change significantly. ...
Full-text available
Ionizing radiation sources such as Solar Energetic Particles and Galactic Cosmic Radiation may cause unexpected errors in imaging and communication systems of satellites in the Space environment, as reported in the previous literature. In this study, the temporal variation of the speckle values on Sentinel 1 satellite images were compared with the cosmic ray intensity/count data, to analyze the effects which may occur in the electromagnetic wave signals or electronic system. Sentinel 1 Synthetic Aperture Radar (SAR) images nearby to the cosmic ray stations and acquired between January 2015 and December 2019 were processed. The median values of the differences between speckle filtered and original image were calculated on Google Earth Engine Platform per month. The monthly median “noise” values were compared with the cosmic ray intensity/count data acquired from the stations. Eight selected stations’ data show that there are significant correlations between cosmic ray intensities and the speckle amounts. The Pearson correlation values vary between 0.62 and 0.78 for the relevant stations.
... The former is the cumulative long term ionizing damage due to protons and electrons, whilst the latter is caused by a single charged particle such as heavy ions and protons [24,25]. The Tenkoh satellite that was developed at Kyushu Institute of Technology and launched in October 2018 was analyzed to have suffered SEEs after passing over the South Atlantic Anomaly [26] . ...
Full-text available
In the recent past, research on the utilization of deep learning algorithms for space applications has been widespread. One of the areas where such algorithms are gaining attention is in spacecraft pose estimation, which is a fundamental requirement in many spacecraft rendezvous and navigation operations. Nevertheless, the application of such algorithms in space operations faces unique challenges compared to application in terrestrial operations. In the latter, they are facilitated by powerful computers, servers, and shared resources, such as cloud services. However, these resources are limited in space environment and spacecrafts. Hence, to take advantage of these algorithms, an on-board inferencing that is power- and cost-effective is required. This paper investigates the use of a hybrid Field Programmable Gate Array (FPGA) and Systems-on-Chip (SoC) device for efficient onboard inferencing of the Convolutional Neural Network (CNN) part of such pose estimation methods. In this study, Xilinx’s Zynq UltraScale+ MPSoC device is used and proposed as an effective onboard-inferencing solution. The performance of the onboard and computer inferencing is compared, and the effectiveness of the hybrid FPGA-CPU architecture is verified. The FPGA-based inference has comparable accuracy to the PC-based inference with an average RMS error difference of less than 0.55. Two CNN models that are based on encoder-decoder architecture have been investigated in this study and three approaches demonstrated for landmarks localization.
Full-text available
In orbit, we find a harsh environment able to damage even space-qualified components. The main threats will be listed in the following lines, one by one, also presenting some of the effects on commercial electronics. According to the literature, the most recommended materials and countermeasures will be also introduced under each 'Materials and Countermeasures' paragraph.
Full-text available
As small satellites are becoming more widespread for new businesses and applications, the development time, failure rate and cost of the spacecraft must be reduced. One of the systems with the highest cost and the most frequent failure in the satellite is the Electrical Power System (EPS). One approach to achieve rapid development times while reducing the cost and failure rate is using scalable modules. We propose a solar module integrated converter (SMIC) and its verification process as a key component for power generation in EPS. SMIC integrates the solar array, its regulators and the telemetry acquisition unit. This paper details the design and verification process of the SMIC and presents the in-orbit results of 12 SMICs used in Ten-Koh satellite, which was developed in less than 1.5 years. The in-orbit data received since the launch reveal that solar module withstands not only the launching environment of H-IIA rocket but also more than 1500 orbits in LEO. The modular approach allowed the design, implementation and qualification of only one module, followed by manufacturing and integration of 12 subsequent flight units. The approach with the solar module can be followed in other components of the EPS such as battery and power regulators.
Full-text available
In addition to traditional interest by various governments in space exploration, there is growing interest in private missions to Mars and other deep space destinations within the next decade. Private missions could consider persons not restricted by radiation limits; however there remains an interest in the level of risk to be encountered. The major risk for space travel is cancer from galactic cosmic rays (GCR), while circulatory diseases in suggested in some but not all epidemiology studies at modest doses (<1 Gy) and detriments in cognition are suggested by rodent studies following acute irradiation with moderate doses of heavy ions. The GCR are not easily shielded since they consist of high energy protons, heavy ions and secondary radiation produced in shielding and tissue. Furthermore heavy ions are more effective per unit dose in causing solid cancers compared to gamma-rays. In addition non-targeted effects (NTEs) are suggested by most low dose radiobiology studies to increase biological effectiveness for low doses of high LET radiation. Astronauts and cosmonauts are typically above 40-y, while younger aged persons could participate in private space missions. In this paper, we describe cancer and circulatory disease risks for a 940 d Mars mission for average solar minimum conditions for persons of varying ages from 20 to 60 years. For the first-time NTEs are considered in Mars mission cancer risk predictions. Cancer morbidity risks and 95% confidence intervals for age 20-y persons are predicted as 20.9% [7.04, 51.4] and 12.7% [4.97, 29.3] for females and males, respectively. We find that cancer fatality risks decline with age of exposure while circulatory disease risks are nearly independent of age of exposure. The ratio of cancer to circulatory disease fatalities decreases from about 8-to-1 at 20 y to 5-to-1 at 60 y in females and 4-to-1 and 2.5-to-1 in males with about 2-times higher loss of life expectancy for cancer deaths compared to circulatory related deaths, indicating the much higher importance of cancer risk compared to circulatory disease risks for persons participating in space missions
Full-text available
The exploration of energization and radiation in geospace (ERG) satellite, nicknamed “Arase,” is the second satellite in a series of small scientific satellites created by the Institute of Space and Astronautical Science of the Japan Aerospace Exploration Agency. It was launched on December 20, 2016, by the Epsilon launch vehicle. The purpose of the ERG project is to investigate how high-energy (over MeV) electrons in the radiation belts surrounding Earth are generated and lost by monitoring the interactions between plasma waves and electrically charged particles. To measure these physical processes in situ, the ERG satellite traverses the heart of the radiation belts. The orbit of the ERG is highly elliptical and varies due to the perturbation force: the apogee altitude is approximately 32,200–32,300 km, and the perigee altitude is 340–440 km. In this study, we introduce the scientific background for this project and four major challenges that need to be addressed to effectively carry out this scientific mission with a small satellite: (1) dealing with harsh environmental conditions in orbit and electromagnetic compatibility issues, (2) spin attitude stabilization and avoiding excitation of the libration by flexible structures, (3) attaining an appropriate balance between the mission requirements and the limited resources of the small satellite, and (4) the adaptation and use of a flexible standardized bus. In this context, we describe the development process and the flight operations for the satellite, which is currently working as designed and obtaining excellent data in its mission.
Full-text available
An optical system for measuring the coefficient of thermal expansion (CTE) of materials has been developed based on electronic speckle interferometry. In this system, the temperature can be varied from -60°C to 180°C with a Peltier device. A specific specimen geometry and an optical arrangement based on the Michelson interferometer are proposed to measure the deformation along two orthogonal axes due to temperature changes. The advantages of the system include its high sensitivity and stability over the whole range of measurement. The experimental setup and approach for estimating the CTE was validated using an Aluminum alloy. Following this validation, the system was applied for characterizing the CTE of carbon fiber reinforced composite (CFRP) laminates. For the unidirectional fiber reinforced composites, the CTE varied with fiber orientation and exhibits anisotropic behavior. By stacking the plies with specific angles and order, the CTE of a specific CFRP was constrained to a low level with minimum variation temperature. The optical system developed in this study can be applied to CTE measurement for engineering and natural materials with high accuracy.
Full-text available
The first major scientific discovery of the Space Age was that the Earth is enshrouded in toroids, or belts, of very high-energy magnetically trapped charged particles. Early observations of the radiation environment clearly indicated that the Van Allen belts could be delineated into an inner zone dominated by high-energy protons and an outer zone dominated by high-energy electrons. The energy distribution, spatial extent and particle species makeup of the Van Allen belts has been subsequently explored by several space missions. Recent observations by the NASA dual-spacecraft Van Allen Probes mission have revealed many novel properties of the radiation belts, especially for electrons at highly relativistic and ultra-relativistic kinetic energies. In this review we summarize the space weather impacts of the radiation belts. We demonstrate that many remarkable features of energetic particle changes are driven by strong solar and solar wind forcings. Recent comprehensive data show broadly and in many ways how high energy particles are accelerated, transported, and lost in the magnetosphere due to interplanetary shock wave interactions, coronal mass ejection impacts, and high-speed solar wind streams. We also discuss how radiation belt particles are intimately tied to other parts of the geospace system through atmosphere, ionosphere, and plasmasphere coupling. The new data have in many ways rewritten the textbooks about the radiation belts as a key space weather threat to human technological systems.
Full-text available
It's a grand opportunity to build new small deep space probe called Shinen2, developed by Kyushu Institute of Technology (KIT), in corporation with the different companies and institutions of engineering in Kagoshima University (Japan), NASA Johnson Space Center, was launched by the rocket H-IIA of Japan Aerospace Exploration Agency (JAXA) with Hayabusa 2, on December 3, 2014 in Tanegashima. This project involves Japanese students and foreigners, permitted a multi-cultural environment and an excellent tools for education. The students are in charge for the design, assembly, integration, tests of the space probe subsystems, and build-up of the existing ground stations facilities for tracking the telemetry data of Shinen2. It will enhance capacity building for the students, and scientific research for upcoming studies. The main approach to carry out the main mission of space probe. In parallel, to develop each subsystem of Shinen2: structure design, system bus architecture including the Communication Control Unit CCU, Power Control Unit PCU specifications, and new Particle Pixel Detector PPD for deep space radiation exploration. The development period for the space probe was only one year; it was extremely a short term. The mass budget and size were strictly limited while requiring a higher reliability. This paper describe a control system design for a small deep space probe which was developed to implement different missions and to satisfy the various requirements listed below.
In 2009 the need for a suitable onboard computer design arose for the small satellite project at the Institute of Space Systems, University of Stuttgart, Germany. It had to meet the constraints imposed by the small satellite (a 130 kg CubeSat) with its full featured ACS, diverse payloads and full CCSDS telecommand and telemetry standard compliance. The design of the Onboard Computer system lead to a functional merging between onboard computer components and the satellite's Power Control and Distribution Unit, resulting in a very innovative solution – the so-called Combined Data and Power Management Infrastructure. The technical implementation of such a design was achieved with the support of an international industry partner consortium consisting of Astrium GmbH, Aeroflex Colorado Springs Inc., 4Links Ltd., Aeroflex Gaisler AB, Vectronic Aerospace GmbH and HEMA Kabeltechnik GmbH & Co. KG. At end of the flight unit's development the consortium decided to provide a single consistent documentation of the developed CDPI Infrastructure. The technical overview should be available for other university students in a sort of mix between technical brochure and user guide. This book also might be of interest for future university or industry partners who intend to order rebuilds / adaptations of the CDPI infrastructure or even the entire satellite bus in Stuttgart for their missions.
The radiation risk radiometer-dosimeter (R3D)-R2 solid-state detector performed radiation measurements at the European Space Agency EXPOSE-R2 platform outside of the Russian “Zvezda” module at the International Space Station (ISS) from 24 October 2014-11 January 2016. The ISS orbital parameters were: average altitude of 415 km and 51.6° inclination. We developed special software and used experimentally-obtained formulas to determine the radiation flux-to-dose ratio from the R3DR2 Liulin-type deposited-energy spectrometer. We provide for the first time simultaneous, long-term estimates of radiation dose external to the ISS for four source categories: (i) galactic cosmic ray particles and their secondary products; (ii) protons in the South Atlantic Anomaly region of the inner radiation belt (IRB); (iii) relativistic electrons and/or bremsstrahlung in the outer radiation belt (ORB); and (iv) solar energetic particle (SEP) events. The latter category is new in this study. Additionally, in this study, secondary particles (SP) resulting from energetic particle interaction with the detector and nearby materials are identified. These are observed continuously at high latitudes. The detected SPs are identified using the same sorting requirements as SEP protons. The IRB protons provide the highest consistent hourly dose, while the ORB electrons and SEPs provide the most extreme hourly doses. SEPs were observed 11 times during the study interval. The R3DR2 data support calculation of average equivalent doses. The 30-day and 1-year average equivalent doses are much smaller than the skin and eyes doses recommendations by the National Council on Radiation Protection (Report 132), which provides radiation protection guidance for Low Earth Orbit.
Weak solar activity and high cosmic-ray flux during the coming solar cycle are qualitatively anticipated by the recent observations that show the decline in the solar activity levels. We predict the cosmic-ray modulation and resultant radiation exposure at flight altitude by using the time-dependent and three dimensional model of the cosmic-ray modulation. Our galactic cosmic ray (GCR) model is based on the variations of the solar wind speed, the strength of the heliospheric magnetic field (HMF), and the tilt angle of the heliospheric current sheet. We reproduce the 22-year variation of the cosmic-ray modulation from 1980 to 2015 taking into account the gradient-curvature drift motion of GCRs. The energy spectra of GCR protons obtained by our model show good agreement with the observations by BESS and PAMELA except for a discrepancy at the solar maximum. Five year annual radiation dose around the solar minimum at the solar cycle 24/25 will be approximately 19% higher than that in the last cycle. This is caused by the charge sign dependence of the cosmic-ray modulation, such as the flat-top profiles in a positive polarity.