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Aerospace 2020, 7, 35; doi:10.3390/aerospace7040035 www.mdpi.com/journal/aerospace
Article
In Situ Measurement of Carbon Fibre/Polyether Ether
Ketone Thermal Expansion in Low Earth Orbit
Farhan Abdullah *, Kei-ichi Okuyama, Isai Fajardo and Naoya Urakami
Department of Applied Science for Integrated Systems Engineering, Kyushu Institute of Technology, 1-1
Sensui, Tobata, Kitakyushu, Fukuoka 804-8550, Japan; okuyama.keiichi008@mail.kyutech.jp (K.O.);
q595902f@mail.kyutech.jp (I.F.); urakami.naoya149@mail.kyutech.jp (N.U.)
* Correspondence: abdullah.farhan835@mail.kyutech.jp
Received: 5 March 2020; Accepted: 24 March 2020; Published: 26 March 2020
Abstract: The low Earth orbit (LEO) environment exposes spacecraft to factors that can degrade the
dimensional stability of the structure. Carbon Fibre/Polyether Ether Ketone (CF/PEEK) can limit
such degradations. However, there are limited in-orbit data on the performance of CF/PEEK. Usage
of small satellite as material science research platform can address such limitations. This paper
discusses the design of a material science experiment termed material mission (MM) onboard Ten-
Koh satellite, which allows in situ measurements of coefficient of thermal expansion (CTE) for
CF/PEEK samples in LEO.
Results from ground tests before launch demonstrated the feasibility of
the MM design. Analysis of in-orbit data indicated that the CTE values exhibit a non-linear
temperature dependence, and there was no shift in CTE values after four months. The acquired in-
orbit data was consistent with previous ground tests and in-orbit data. The MM experiment
provides data to verify the ground test of CF/PEEK performance in LEO. MM also proved the
potential of small satellite as a platform for conducting meaningful material science experiments.
Keywords: atomic oxygen; coefficient of thermal expansion; carbon fibre; polyether ether ketone
(PEEK); low Earth orbit; ultraviolet; in situ; small satellite; thermal cycle; in-orbit
1. Introduction
The Low Earth Orbit (LEO) environment exposes spacecraft to ultraviolet (UV) radiation, atomic
oxygen (AO), vacuum and temperature variation. These factors can affect the dimensional stability
of a spacecraft structure.
Dimensional stability is crucial for different spacecraft structures especially parts that deal with
precision such as antenna, truss structure and optical support structure. For example, minute
dimensional changes can result in a serious loss of signal for an antenna due to loss in pointing
accuracy. Therefore, it is beneficial to utilize dimensionally stable material in manufacturing a
spacecraft part. However, the high cost of launching payloads to space place a premium on spacecraft
mass. Therefore, it is desirable to use a lightweight and dimensionally stable material for spacecraft
structures. Composite materials such as polymer matrix composites (PMC) can suit the mentioned
requirements partly due to high strength to weight ratio, high stiffness and low coefficient of thermal
expansion (CTE) [1,2].
In the case of PMC, the main factors affecting the dimensional stability are moisture, thermal
expansion, mechanical loading and microyielding [3]. Thermal expansion is caused by repeated
thermal cycling due to the temperature variation when a satellite passes from direct sunlight into
Earth shadow. Microyielding is caused by microcracking in the PMC. Repeated thermal cycling can
induce microcracking [4]. These microdamage develops because of stresses caused by the fibre-
matrix CTE mismatch, the CTE mismatch in properties along and transverse to the fibre direction
and through the ply or lamina [5]. The microcracks can increase as the number of thermal cycles
Aerospace 2020, 7, 35 2 of 24
increased. Moreover, the properties of the fibre and matrix can affect the extent of microcracking.
This process can lead to a progressive change in the CTE because thermal expansion is affected by
microcracking behaviour [4,6,7].
Application of high-performance PMC in spacecraft structures is crucial to limit changes to
dimensional stability. Carbon Fibre/Polyether Ether Ketone (CF/PEEK) composite is a high-
performance PMC due to its inherent dimensionally stable properties. PEEK is a semi-crystalline
thermoplastic polymer. The toughness of PEEK resin provides excellent resistance to microcracking
induced by thermal cycling [4,8]. Previously, it was mentioned that microcrack develops due to
internal stress caused by fibre-matrix CTE mismatch. In a thermoplastic composite such as CF/PEEK,
internal stress is dissipated internally within the structure instead of through microcracking. Heat is
generated due to the internal dissipation [4]. As a result, microcracking can be minimized and
changes to CTE kept to a minimal level.
In a previous satellite project by Kyushu Institute of Technology (KIT), CF/PEEK was used as
the primary material for the external structure of Shinen-2 deep space probe [9]. The structure
survived the launch and space environment. However, more space data is required to further
understand the behaviour of CF/PEEK including dimensional stability in space. Previous sources of
CF/PEEK data mainly originated from ground test [4,10,11].
The lack of space performance data can be attributed to various factors. Firstly, there is limited
access to space to launch satellites or experiments into orbit. The current launch options are initially
as cargo to International Space Station (ISS), then deployment from the Japanese Experiment Module
on ISS (KIBO), as a secondary payload on a rocket, launch in a cluster with other small satellites and
as a primary payload on a dedicated small launch vehicle [12]. The ISS cargo and secondary payload
placed a constraint on the type of orbit for the payload. Even though numerous companies are
developing small launch vehicles, most of the developments may not reach maturity thus limiting
access to space. Currently, Northrop Grumman Pegasus, Japan Aerospace Exploration Agency
(JAXA) Epsilon and Rocket Lab Electron rocket are the only operational small launch vehicles [13,14].
Secondly, there are limited retrieval options to retrieve samples from orbit. The European Space
Agency (ESA) Space Rider, Soyuz capsule and SpaceX Dragon capsule allow the option to retrieve
samples from space but are expensive to operate and have limited flight frequency [15–17]. Thirdly,
ground test offers a lower testing cost; however, the challenge of simulating actual space environment
tends to decrease the accuracy of results [18,19].
The emerging small satellite market provides a promising solution to the lack of in-orbit data
[20,21]. The small satellite provides an available platform for space experiments including material
science experiments [22]. In particular, CubeSats are a promising orbital research platform due to low
development cost and accessible to a wider group of participants [23]. The miniaturization of
components such as lab-on-a-chip (LOC) and microelectromechanical systems (MEMS) allows the
creation of smaller hardware that can fit into smaller satellites [15].
There is a variety of techniques to measure the dimensional stability of CF/PEEK. This paper
focuses on measuring the thermal expansion factor. This paper will discuss the material science
experiment termed as material mission (MM) and is one of several payloads onboard the Ten-Koh
satellite shown in Figure 1. Ten-Koh satellite was developed by KIT and was successfully launched
on 29 October 2018. Ten-Koh satellite orbits the Earth in a sun-synchronous sub-recurrent orbit at an
altitude of approximately 600 km. The purpose of the MM is to perform in situ measurements of
changes in the CTE of CF/PEEK composites samples in LEO. Strain gauges and temperature sensors
were used to measure changes in strain and temperature for calculation of CTE. The experiment
eliminates the need for sample retrieval by transmitting results to the ground station. This paper
briefly introduces the system architecture of the MM design. Ground validation test and in-orbit data
were also presented and compared. Discussions on the ground and in-orbit data are provided
together with issues and improvements for MM.
Aerospace 2020, 7, 35 3 of 24
(a)
(b)
Figure 1. Ten-Koh satellite configuration and location of material mission (MM). (a) Ten-Koh flight
model (FM); (b) location of MM on top of Ten-Koh external structure.
2. Materials and Methods
2.1. Material Mission System Architecture
The MM consisted of two main components; the internal printed circuit board (PCB) and the
external PCB. The external PCB contains the CF/PEEK samples, strain gauges and temperature
sensors. The external PCB was installed on the top plate of the external satellite structure for
maximum field of view as shown in Figure 1. The internal PCB contains the MM electronic circuit for
operating the experiment. Figure 2 illustrates the internal and external PCB, respectively. Both PCB
were in the upper section of the Ten-Koh satellite.
Aerospace 2020, 7, 35 4 of 24
(a) (b)
Figure 2. MM main components. (a) Internal printed circuit board (PCB) contains MM electronic
circuit; (b) external PCB contains the Carbon Fibre/Polyether Ether Ketone (CF/PEEK) samples, strain
gauges and temperature sensors.
MM was designed to consider the following three constrains:
1. The experiment should be able to survive launch conditions without being damaged.
2. The experiment will not generate unintentional space debris.
3. The limited development time due to the piggyback nature of this mission.
For the first reason, the samples and the experiment assembly can be destroyed and scattered
due to the harsh launch environment leading to possible damage of the main satellite and other
satellites in the rocket fairing. Therefore, usage of mechanical devices or moving parts to create strain
on the samples may fail or cause damage during the launch environment. For the second reason,
prolonged exposure to space can deteriorate the mechanical properties of the samples. The samples
under due stress can be damaged or break apart leading to an increase in space debris. The second
reason was the primary factor for omitting destructive tests such as tensile test. The reason is that the
destruction of the samples can create space debris. Moreover, when the strain gauge is attached to
the surface of the specimen using an adhesive, the adhesive may outgas in the vacuum environment.
The adhesive and detached strain gauge can add to the existing space debris in LEO.
Due to the aforementioned constrains, a passive measurement system utilizing non-destructive
test was preferred over an active system. Measurement of CTE due to thermal strain was selected
because it requires only a passive measuring system. According to the thermal simulation results
performed before launch, the temperature range of the external structure is estimated to be between
10 and 40 °C, which can provide sufficient change in thermal strain.
The dimensions of the external PCB measured 78 × 80 mm. Three CF/PEEK samples were bolted
to the external PCB. Each sample consisted of two pieces of CF/PEEK thermally welded together
using a heat press machine. The overall dimension of each sample was 50 mm long, 10 mm wide and
2 mm thick, as illustrated in Figure 3.
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Figure 3. Detail drawing of CF/PEEK samples measuring 50 mm long, 10 mm wide and 2 mm thick.
The CF was made from plain-woven carbon fabric manufactured by Toray with a 0/90° pattern.
The PEEK resin was manufactured by Victrex (Lancashire, United Kingdom). The main reason for
the plain-woven pattern was to maintain a quasi-isotropic property for the external structure and the
ease of woven material to conform to the moulding tool with complex shape [24]. A protective coating
was applied to two of the three samples as shown in Table 1. The purpose of the protective coating
was to determine the changes in CTE due to different factors in space.
Table 1. List of protective coating applied to MM CF/PEEK samples.
Degradation Factor Type of Coating
Atomic oxygen Silsesquioxane (RSiO
3/2
)
Ultraviolet radiation Yttrium oxide (Y
2
O
3
)
Each sample contains a single 0°/90° 2-element rosette stacked type strain gauge. Carbon Fibre
Reinforced Thermoplastic (CFRTP) melts when heated above the melting temperature [25]. This
unique feature allows a strain gauge to be attached between two pieces of CF/PEEK by thermal
welding as shown in Figure 4. The rationale for thermally welding the strain gauge was to avoid
accidental detachment of the strain gauge in orbit. Outgassing of the standard strain gauge adhesive
can cause the strain gauge to become detached from the samples in orbit. Initially, a strain gauge was
placed on one piece of CF/PEEK. Each element of the strain gauge was aligned to the fibre direction.
Another piece of CF/PEEK was placed on the strain gauge and the first piece of CF/PEEK. Both pieces
with the strain gauge in between were wrapped in aluminium tape and placed in a hot press machine
(FT-10HP, Full Tech, Japan). The temperature was increased stepwise until a maximum of 400 ˚C.
Moisture was removed from the sample by maintaining a constant temperature for approximately 30
min after each temperature increment. Pressure was not applied to the samples to avoid damaging
the lead wire. Instead, the sample was placed on one hot press plate with the other plate located near
and above the sample.
(a) (b)
Figure 4. Strain gauge assembly. (a) Exploded view of a MM sample with the strain gauge thermally
welded between two pieces of CF/PEEK; (b) MM sample after completion of thermal welding process.
Aerospace 2020, 7, 35 6 of 24
The measurement of sample temperature by attachment of thermocouple to sample is not viable
due to the gradual outgassing of adhesive bonding in a vacuum environment. Therefore, MM applied
a temperature measurement system that utilized heat conduction. The system is illustrated in
Figure 5. Temperature measurement for each sample was measured using an AD590 temperature
sensor (AD590, Analog Devices, Norwood, MA, USA) attached underneath the external PCB but
below each sample. The AD590 sensor is capable of recording temperatures between −55 and 150 °C
with an error range of ±1 ˚C. Heat was conducted via conduction from the surface of the sample
through a hole filled with thermal grease and α-gel to the temperature sensor.
Figure 5. Cutaway of MM sample and temperature measurement system. The strain gauge is
thermally welded between two CF/PEEK samples. Temperature is measured using a temperature
sensor via heat conduction from the surface of the sample through a hole filled with thermal grease
and α-gel.
2.1.1. Electrical Design
The MM was developed to measure the CTE of CF/PEEK. CTE is the change in length or volume
of a material as a function of temperature rise. The calculation of CTE is based on Equation (1) [1].
= −
1
−
=∆
1
∆
(1)
where α is CTE (ppm/°C), L is length (m), L
0
is initial length (m), T is temperature (°C), T
0
is initial
temperature (°C), ∆L is change in length (m) and ∆T is change in temperature (°C). The values of CTE
are not always constant. Therefore, the CTE is expressed as an average value in a certain temperature
range [1].
The MM has two main sensors to calculate CTE; strain gauge and temperature sensor. Each
sample has one strain gauge and a temperature sensor attached to it, as shown in Figure 5.
2.1.2. Strain Gauge
Strain is the deformation of a material due to a force or a set of forces resulting in a change in
length as shown in Equation (2) [26]. In the context of MM, strain is the change in length of the
CF/PEEK sample due to temperature. A strain gauge detects this strain as electrical signals. Strain
gauge electrical resistance changes with deformation. When a strain gauge is bonded to a material,
the deformation of the material changes the electrical resistance of the strain gauge. The change in
strain is small and is measured in microstrain (με). The relation between strain and change in
resistance is shown in Equation (3) [26].
= ∆
(2)
where ε is strain (μm/m), L is the initial length (m) and ∆L is the elongation.
Aerospace 2020, 7, 35 7 of 24
∆
=
(3)
where R is resistance (Ω), ∆R is change in resistance (Ω), Ks is gauge factor and ε is strain (μm/m).
Gauge factor is the strain gauge sensitivity. A Wheatstone bridge circuit is commonly used to convert
minute change in resistance to voltage change (∆V) or output voltage (e
0
). MM utilized a quarter
bridge type of the Wheatstone bridge circuit. The e
0
was calculated using Equation (4) [26].
0
=1
4∆
= 1
4
(4)
where e
0
is output voltage (V), E is bridge excitation voltage or input voltage (V), R is resistance (Ω),
∆R is change in resistance (Ω), K
S
is gauge factor and ε is strain (μm/m). Based on Equation (4) strain
is proportional to the output voltage.
In the LEO environment, the temperature outside of Ten-Koh regularly changes between hot
and cold due to the orbital movement around Earth. The apparent strain was generated by the strain
gauge bonded to the CF/PEEK sample due to the varying temperatures. The formula for the apparent
strain is shown in Equation (5) [26]. The sources of the apparent strains are listed below:
1. CTE mismatch between resistance element and measure material.
2. Change in resistance of gauge with temperature or resistivity of strain gauge grid.
=
+
−
∆
(5)
where ε
T
is the apparent strain (μm/m), β is resistive temperature coefficient of resistive element, K
s
is gauge factor, α
m
is CTE of measured material, α
g
is CTE of strain gauge and ∆T is change in
temperature (°C). Apparent strain can be reduced if α
m
and α
g
are nearly equal. However, for MM,
the challenge of selecting a matching CTE between measured material and strain gauge may result
in inaccurate measurement due to apparent strain. Therefore, an alternative apparent strain removal
method was selected for MM as shown in Figure 6 [27]. The two white holes in Figure 6 represent the
placement of the bolts on the MM samples. In the horizontal direction or direction 1, the sample was
constrained with bolts. An arbitrary temperature increase was applied resulting in compressive strain
as shown in Equation (6). In the vertical direction or direction 2, the same temperature change as
above was applied. However, in this direction, the sample was unconstrained resulting in tensile
strain as shown in Equation (7). The same amount of apparent strain included in both directions can
be offset by taking the difference between strain measurement in direction 1 and 2 as shown in
Equation (8). The accuracy of the strain measurement was increased by employing the mentioned
method. The selection of a single 0°/90° 2-element rosette stacked strain gauge was due to the
requirement of the apparent strain removal method.
Figure 6. Alternative apparent strain removal method applied by MM. The difference between strain
measurement in direction 1 and 2 can offset the amount of apparent strain included in both directions.
1:
= −
+
(6)
where ε
1
is strain in direction 1, ε
m
is strain of sample, and ε
g
is strain of strain gauge
2:
=
+
(7)
Aerospace 2020, 7, 35 8 of 24
where ε
2
is strain in direction 2, ε
m
is strain of sample, and ε
g
is strain of strain gauge
|
−
|= 2
(8)
The block diagram for the strain gauge circuit is shown in Figure 7. The ADR4520 voltage
regulator (Analog Devices, Norwood, MA, USA) was used to convert the 5 V power supply from
battery to 2 V input voltage for the strain gauge. The low noise (0.1 Hz to 10 Hz) and wide operating
temperature (−40 to 125 °C) were the reasons for selecting ADR4520. The LM358-N (Texas
Instruments, Dallas, TX, USA) operational amplifier was used to provide high input impedance. The
reason for the high input impedance was to minimize the drop in the voltage input. The LM358-N
was selected due to its proven flight heritage in the Shinen-2 project. An IRLML6244 MOSFET
(Infineon Technologies, Neubiberg, Germany) was used to provide a switching circuit. The varying
temperature in LEO can also affect the resistance of the lead wire. The 3-lead wire strain gauge was
selected for self-temperature compensation. The minute change in resistance was converted by the
Wheatstone quarter bridge to voltage change, ∆e
0
. The ADS1220 (Texas Instruments, Dallas, TX,
USA), a 24-bit analogue-to-digital converter (ADC) was used to amplify the value of ∆e
0
which was
in the μV range. Each of the 2-element strain gauges was connected to two ADC with one ADC for
each gauge element.
Figure 7. Strain gauge circuit block diagram.
2.2. Mission Operations
The general process flow for MM starts when the ground station transmits a command to the
internal PCB via onboard computer (OBC) to read temperature and strain measurements. Data is
then transmitted to the ground station. A decoder program developed using C language converts the
data packets into strain, temperature and CTE values. Figure 8 illustrates the overall MM mission
operations process flow. The user can set the number of measurements or packet number for each
operation session. One packet of data has a size of 60 bytes. The usage of amateur radio with limited
bandwidth for communication between the ground station and satellite limits the amount of MM
data that can be downloaded in one pass. Moreover, Ten-Koh has other payloads that generate data.
This also contributes to limit the amount of MM data that can be transmitted to the ground station.
As a result, MM operations were limited to real-time measurements when Ten-Koh passes over the
KIT ground station.
Aerospace 2020, 7, 35 9 of 24
Figure 8. Overview of MM mission operations process flow.
2.3. Ultraviolet Sensor
Originally, there were two UV sensors, one sensor can detect all UV wavelengths and the other
one can detect only UV-C. UV filters were proposed for filtering out UV-A and UV-B, thus allowing
the UV sensor to focus on measuring UV-C radiation. However, the UV sensors were excluded from
the FM due to a flight qualification problem. Further discussion on the UV sensors will be covered in
the discussion section.
2.4. Ground Validation Test
A ground validation test was performed to determine the effectiveness of MM measurement
system. The test was carried out in a thermal vacuum chamber (TVAC) in KIT. The TVAC can
produce vacuum conditions up to 1.0 × 10
−5
Pa. The purpose of the test is listed as follows:
1. To determine the effectiveness of the temperature measurement method using the AD590
temperature sensor.
2. To determine the effectiveness of the CTE measurement system.
The surface temperature of the sample recorded by thermocouple was compared with those
measured by the temperature sensors. The CTE measurement by the MM circuit was compared with
CTE data obtained using a TM-7000 (Advance Riko, Yokohama, Japan) Thermomechanical Analyzer
(TMA) provided by Agne Technology Center Company, Tokyo, Japan. Identical external PCB,
internal PCB and CF/PEEK samples designed for the FM model were used for the ground test. The
number of samples was three which was the same with the FM model. Thermocouples were attached
to the sample and temperature sensors for comparison purposes. Polyimide heaters were attached to
the heater frame made of aluminium to allow uniform distribution of heat to the samples. Heat was
indirectly distributed to the samples via radiation from the aluminium frame. Black paint was also
applied to the aluminium frame to increase the amount of radiated heat. The test setup is illustrated
in Figure 9.
Figure 9. Full system test setup. Identical external and internal MM PCB to flight model were used
for the test. Polyimide heaters attached to aluminium frame were used as a heat source.
Aerospace 2020, 7, 35 10 of 24
The pressure in the TVAC was set to below 1.3 × 10
−3
Pa. The range of temperature change was
between −40 and 22 °C. This temperature range was selected based on recorded temperature obtained
from space. The test was carried out in two scenarios. The first was to cool down the test chamber
from a room temperature of 22 to −40 °C at a cooling rate of −0.5 °C/min. In the second scenario, the
test chamber temperature was increased from −40 to 22 °C at a rate of 1 °C/min.
3. Results
The function of MM was to perform in situ measurement of CTE of CF/PEEK samples in LEO.
Ground validation test was performed to validate the design of MM. In-orbit data validated the
feasibility of MM to provide in-orbit measurement of CTE in LEO. MM used strain gauges and
temperature sensors to measure strain and temperature for calculating CTE. CTE is the change in
length of material as a function of temperature rise or fall. CTE is calculated based on Equation (1).
3.1. Ground Validation Test Results
Table 2 lists the maximum, minimum and median values for the difference between temperature
measured by temperature sensor and thermocouple in each phase for each sample. Sample 1 refers
to the sample with no protective coating; sample 2 refers to the sample with UV protective coating,
and sample 3 refers to the sample with AO protective coating. Median was included to provide an
accurate assessment of the difference. In the case of sample 3, the median difference was within
±1 °C. This was within the measurement error range for AD590 which is ±1 °C as stated in the
datasheet [28]. Sample 1 showed the same median deviation range except during the increasing
phase. However, for sample 2, the median deviation range is up to 2.67 °C during the increasing
phase. The MM temperature measurement method used an indirect method to measure heat.
Therefore, there was additional heat input from the external PCB that has a minor effect on the
temperature readings. In terms of the overall trend, there is no adverse difference between the
temperatures recorded by thermocouples and temperature sensors. The MM temperature
measurement method is feasible based on the obtained results.
Table 2. Maximum, minimum and median value for the difference between temperature measured
by temperature sensor and thermocouple in each phase for each sample.
Test Phase
Sample 1 Sample 2 Sample 3
Max.
1
(°C)
Min.
2
(°C)
Median
(°C)
Max.
1
(°C)
Min.
2
(°C)
Median
(°C)
Max.
1
(°C)
Min.
2
(°C)
Median
(°C)
Equilibrium 0.71 −0.31 0.43 2.29 1.18 1.89 0.55 −0.56 0.10
Increasing
Temperature 2.36 0.09 1.45 3.72 0.40 2.67 2.00 −0.71 0.68
Decreasing
Temperature 2.22 −2.30 −0.07 3.90 −0.79 1.59 1.32 −2.96 −0.64
1
Max. refers to maximum.
2
Min. refers to minimum.
The overall results from the ground test are shown in Figure 10. For each data point, the median
values of strain and temperature range were used for calculation of CTE. Figure 10 showed that the
value of CTE varied with increasing temperature. This was consistent with results from previous
studies [4,8,29,30]. Sample 2 had the highest range of CTE compared to sample 1 and 3. CTE varied
between 2.98 and 3.99 ppm/˚C for sample 1 and between 2.93 and 3.23 ppm/˚C for sample 2 for
temperature up to 10 ˚C. However, after 10 ˚C there was a big increase in CTE value up to a maximum
of 8.37 ppm/˚C and 9.95 ppm/˚C for samples 1 and 2, respectively. For sample 3, CTE remained stable
until 17 ˚C before increasing to 4.15 ppm/˚C. The difference in CTE values between all samples was
due to misalignment between strain gauge elements and the sample fibre direction. A comparison
was performed between the median TMA value of 3.11 ppm/˚C and all the samples. Sample 1 had
Aerospace 2020, 7, 35 11 of 24
the smallest difference between median CTE value and the TMA value, which was 0.14 ppm/˚C.
Sample 2 had the largest difference between median CTE value and the TMA value which was
0.74 ppm/˚C. Overall, the measurement error range between the median TMA value and sample
values was approximately within the range of ±1.00 ppm/°C. The minimal error can be considered
within measurement tolerance. The MM CTE measurement system was proven to be feasible based
on the given results.
Figure 10. The overall result from the full system test. The CTE values for samples 1, 2 and 3 were
compared with reference CTE data measured using Thermomechanical Analyzer (TMA).
3.2. In-orbit Results
MM operations were performed from 9 November 2018 to 19 March 2019. A total of 101
measurements were performed during the mentioned period. The subsequent sections discussed the
results based on the collected measurements.
3.2.1. Temperature Measurement
Figure 11 illustrates the variation of temperature against time. The temperature varied
approximately between −40 and 20 °C for each orbit. This was within the temperature sensor
measurement range of −55 and 150 °C [28]. All samples displayed approximately similar temperature
readings. MM sessions operated during different times for each session. This explains the difference
in temperature reading for each session as shown in Figure 11. Temperatures were generally positive
during daylight pass and negative during night-time pass. Occasionally, the temperature was
negative during daylight pass, which was due to MM external PCB facing away from sunlight.
A comparison was performed with measured temperature value from temperature sensors on
six solar panels nearest to the MM. Each solar panel has a temperature sensor like the one used for
MM. The location of six solar panels near MM is shown in Figure 12. The purpose of the comparison
was to validate the MM temperature sensor reading. The measured temperature from MM was
within the range of the upper and lower limit from the solar panels, which was 50 and −45 °C,
respectively. It should be noted there were minor variations between MM temperature sensor and
solar panels. The temperature sensor of a solar panel was attached to aluminium. The thermal
Aerospace 2020, 7, 35 12 of 24
conductivity of aluminium is higher compared to CF/PEEK which contributed to the variation. Based
on measured data and comparison with solar panels, the MM temperature sensor provided a feasible
method to measure temperature.
Figure 11. Variation of temperature measurement for each sample against time. The temperature
varied between −40 and 20 ˚C in orbit.
Figure 12. Location of six solar panels (SP) near to MM.
3.2.2. CTE Measurement
The changes in CTE against temperature for samples 1, 2 and 3 are shown in Figure 13.
Sample 1 refers to the sample without a protective coating; sample 2 refers to the sample with AO
protective coating, and sample 3 refers to the sample with UV protective coating. The CTE showed a
non-linear temperature dependence, which was consistent with results from previous studies
Aerospace 2020, 7, 35 13 of 24
[4,8,29,30]. Each data point represents the median CTE value calculated based on a range of
temperature and strain measurement during a mission operation session. Moreover, each data point
represented measurement from different date or different time within the same day. Each operation
session has a varying degree of heating or cooling rate due to the absence of an active attitude control
system (AACS) in Ten-Koh. The variation in heating or cooling rate resulted in the scattered CTE
values [31,32]. Further discussion on the effect of variation in heating or cooling rate on the CTE
values will be provided in the subsequent section. The curves for each sample consisted of scattered
data points but follow a consistent upward trend. A bisquare method was used to fit a second-order
polynomial curve due to the degree of distribution. The curve fit expressions for each curve were
provided in the caption of Figure 13. Figure 13 also showed a high concentration of CTE values
between −10 and 10 °C. Limited satellite pass time, communication bandwidth and operation time
due to other onboard experiments placed a constraint on material mission operation duration.
Therefore, the average number of measurement sessions was approximately one operation per day,
which limits the potential to measure CTE values at a wider range of temperatures. Comparison of
CTE between samples shows a varying CTE curve. Sample 3 had the highest CTE values followed
by sample 1 and 2. The difference in CTE values between all samples was due to a slight strain gauge
misalignment with respect to the fibre direction. The strain gauge misalignment will be further
discussed in the subsequent section. Table 3 shows the average change in CTE per ˚C. Between
−10 and 10 °C, all sample curves showed a gradual variation in CTE with temperature between
0.064 and 0.094 CTE/˚C. However, the variation in CTE sharply increased after 10 °C with a maximum
value of 1.018 CTE/˚C for sample 2.
Figure 13. Comparison of CTE measurements between samples 1, 2 and 3. Curve fit for sample 1:
= 0.005052
+0.07656+ 1.349, curve fit for sample 2: = 0.00464
+0.1428+ 3.408 and
curve fit for sample 3: = 0.002033
+0.09219+ 2.348.
Table 3. Average change in CTE with increasing temperature.
Sample Average CTE Change (CTE/˚C)
−10 to 10 °C >10 °C
1 0.064 0.550
2 0.169 1.018
Aerospace 2020, 7, 35 14 of 24
3
0.094
0.771
In terms of long-duration observation, there was no shift in CTE values for up to 4 months or
120 days as shown in Figure 14. The CTE variation remains reasonably similar up to 120 days. There
was no upward or downward shift in the CTE curve. The result is in line with previous studies from
the National Aeronautics and Space Administration (NASA) Long Duration Exposure Facility
(LDEF) mission involving CF/Epoxy samples [33]. However, there were no additional in-orbit data
after 4.5 months due to a loss of communication between Ten-Koh and KIT ground station on
19 March 2019. The likely reason for Ten-Koh failure was due to radiation damage triggered by a
single event effect [34]. Currently, recovery operations are still trying to recover normal Ten-Koh
functions. Nevertheless, these results provide compelling evidence that the MM experiment can
provide reasonably accurate CTE values in LEO for CF/PEEK samples.
Figure 14. Monthly CTE measurement. (a) Sample with no coating; (b) sample with atomic oxygen
(AO) coating; (c) sample with ultraviolet (UV) coating.
4. Discussion
In this study, the development and operations of MM were presented. A ground test was
performed to validate the design of MM. In-orbit data were presented to illustrate the feasibility of
MM to provide in situ measurement of CTE in LEO. The full system test proved the feasibility of the
MM CTE measurement system including the strain gauge circuit and temperature measurement
method.
In-orbit temperature data was comparable with temperature data from nearby solar panels. CTE
variation with temperature was consistent with results from previous studies. Based on Figure 13,
the plot shows an upward curve pattern. In a CF/PEEK composite, the CF has a lower CTE compared
to PEEK or the matrix [8]. Based on Figure 13, all samples exhibit a similar upward curve. At
Aerospace 2020, 7, 35 15 of 24
temperature below 10 °C, CTE is lower and almost constant compared to above 10 °C. At lower
temperature, the shrinking matrix is constrained by the fibres. Thus, fibres are dominant in lower
temperature. As the temperature rises above 0 °C, the CTE values begin to increase in a non-linear
pattern. From this, it might be inferred that the matrix is gradually more dominant in affecting the
CTE values of CF/PEEK.
Figure 13 showed scattered CTE points but with a consistent upward curve. As mentioned in
Section 3.2.2, the difference in heating and cooling rate in LEO affected the change in CTE. This
resulted in the scattered points for CTE values. The effect of the variable heating and cooling rate on
CTE had been shown in previous studies [31,32]. In the case of Ten-Koh, the variation in the heating
and cooling rates was due to the absence of an AACS and few other variables listed below:
1. Phases of Ten-Koh’s orbit, e.g., in the eclipse or the sunlit phase.
2. The orientation of MM external PCB with respect to the sun.
3. Ten-Koh’s rate of rotation.
4. Ten-Koh’s direction of rotation.
The above variables affected the heating and cooling rate of MM samples through the amount
of change in temperature and time. Figure 15 showed the variation in heating and cooling rate for the
sample with UV coating between December 2018 and March 2019. The heating rate varied between
0.39 and 3.24 ˚C/minute while the cooling rate varied between −0.20 and −1.60 ˚C/minute. The cooling
phase occurred mainly during Ten-Koh’s late-night passes over KIT ground station when Ten-Koh
was in Earth’s shadow. The heating phase occurred during afternoon passes when Ten-Koh was in
the sunlit phase in orbit.
Figure 15. Heating and cooling rates for MM samples between December 2018 and March 2019. (a)
Heating rate; (b) cooling rate.
Further analysis was performed to compare results from the ground validation test and in-orbit
data. Figure 16 showed a comparison between ground validation test and in-orbit data. All plots in
Figure 16 showed minor changes or near stable CTE values between −10 to 10 ˚C for both ground
and in-orbit data. The change in CTE increased sharply after 10˚C for ground and in-orbit data except
for sample with AO coating. This sample showed increased CTE at a later temperature. Table 4
shows the average change in CTE with temperature for ground and in-orbit data. However, based
on Table 4, there was difference in CTE value especially in the region below 10 ˚C for the sample with
no coating and sample with UV coating.
Aerospace 2020, 7, 35 16 of 24
Figure 16. Comparison of CTE values between ground validation test data and in-orbit data for all
MM samples. (a) Sample with no coating; (b) sample with AO coating; (c) sample with UV coating.
Table 4. Comparison of the average change in CTE with temperature for ground and in-orbit data.
Sample Type Data Source Average CTE Change (CTE/˚C)
−10 to 10 °C >10 °C
No coating Ground 0.021 0.59
In-orbit
0.064
0.55
AO coating Ground 0.013 0.16
In-orbit
0.170
1.02
UV coating Ground 0.012 0.74
In-orbit
0.094
0.77
Table 5 showed the difference in CTE between ground and in-orbit data for all MM samples.
The sample with no coating had the largest CTE difference while the sample with AO coating had
the smallest CTE difference for temperatures below 10 ˚C. In the temperature range above 10 ˚C, the
sample with AO coating had the largest CTE difference while other samples showed nearly similar
CTE values between ground and in-orbit data.
Table 5. Comparison of the difference in CTE between ground and in-orbit data.
Sample Type
−10 to 10 °C
>10 °C
No coating
1.94
0.60
AO coating 1.00 3.37
UV coating 1.53 0.82
The difference in heating and cooling rate between ground test conditions and LEO environment
caused the difference in the average change in CTE with temperature. The heating and cooling rates
for the ground validation test were 1.0 ˚C/minute and −0.5 ˚C/minute, respectively. The ground
Aerospace 2020, 7, 35 17 of 24
validation test was performed in a single day with the same heating and cooling rate. The ground
validation test was focused on testing the feasibility of the MM measurement system. The test did
not accurately simulate the actual conditions in the LEO environment and the absence of an AACS
in Ten-Koh. This includes the variation in heating and cooling rate in-orbit.
The difference in CTE as shown in Table 5 between ground test and in-orbit data was caused by
the misalignment of strain gauge with respect to the fibre direction of a MM sample. The strain gauge
misalignment also caused the minor difference in CTE between MM samples in-orbit as shown in
Figure 13. A strain gauge misalignment test was performed to compare the CTE values for different
offset positions of the strain gauge with respect to the fibre direction. As a result, the change in CTE
with temperature varied between ground test and in-orbit data.
Strain measurements from three identical CF/PEEK samples were compared using a hot press
machine (FT-10HP, Full Tech, Japan). The CF/PEEK samples have similar material properties but
different dimensions compared to the samples flown to orbit. The sample dimension was 50 mm
long, 10 mm wide and 1 mm thick. A strain gauge was attached to each sample using adhesive to
provide strain measurements. Each strain gauge is a single 0°/90° 2-element rosette stacked type
strain gauge. There were three test scenarios with each scenario representing a different orientation
of the strain gauge with respect to the fibre direction. The strain gauge orientation for each scenario
is shown in Figure 17. The rationale behind the differing orientation was to observe changes in CTE
variation with different offset positions. A thermocouple was attached to the top of each sample for
temperature measurement. Figure 18 shows the test assembly for the offset test. In each scenario,
samples were placed on the lower part and enclosed with a 5 mm-thick metal jig to maintain a
constant heat on the sample from the heat press. The thermocouple was connected to a data logger,
and the strain gauge was connected to a dynamic strainmeter to record data during the test. Test
temperature was varied between 30 and 50 °C to stimulate expansion to the CF/PEEK sample. The
heating rate was approximately 1 °C/min.
Figure 17. Strain gauge orientation with respect to fibre direction for three different scenarios.
Aerospace 2020, 7, 35 18 of 24
Figure 18. Test assembly for strain gauge offset test.
Figure 19 showed changes in CTE due to variation in the orientation of the strain gauge with
respect to the fibre direction. In the first scenario, the strain gauge was aligned to the 90° and 0° fibre
direction. This scenario was the reference for CTE comparison. Figure 19a showed the CTE for strain
gauge element 1 aligned to 90° fibre direction. Figure 19b showed the CTE readings for strain gauge
element 2 initially aligned to 0° fibre direction. The direction of the shift in CTE was shown by the
dotted red arrow in Figure 19a,b. In Figure 19a, scenario 3 exhibited the largest shift in CTE to
negative. Moreover, the shift in CTE towards negative increased as the angle between element 1 and
90° fibre direction increased from 0° to 60°. The same pattern was shown in Figure 19b for the angle
between element 2 and 0° fibre direction albeit in a lower shift increment. It is observed that the same
pattern appeared in the in-orbit data shown in Figure 13 with different CTE for each sample and in
Figure 16 for CTE comparison between ground and in-orbit data. Therefore, misalignment of strain
gauge with respect to fibre direction caused the difference in CTE for different MM samples and
comparison between ground and in-orbit data.
Aerospace 2020, 7, 35 19 of 24
Figure 19. Changes in CTE due to various offset positions. (a) Changes in CTE for element 1 of strain
gauge.; (b) Changes in CTE for element 2 of strain gauge.
There was no shift in CTE values for up to four months as shown in Figure 14. The result is in
line with previous studies. However, in-orbit data was limited to four and a half months due to loss
of data reception from Ten-Koh. A previou s study showed that 100 thermal cycles fr om
−160 to +120 °C produced a minor change in CTE of CF/PEEK composites, the main reason being the
tough property of PEEK matrix [4]. Previously in 1984, NASA conducted the LDEF mission. Several
materials were exposed to the LEO environment including CF/Epoxy samples. After 371 days in
space, there was no substantial degradation in CTE value compared to pre-launch CTE values [33].
Selected results are shown in Table 6 below. As mentioned in the Introduction, microcracking
induced by thermal cycling can affect CTE [4,6,7]. Thermal cycle can be considered as low-cycle
thermal fatigue [5]. In a thermoset composite such as CF/Epoxy, internal stress due to fatigue causes
internal cracks. However, the internal stress is absorbed by the internal structure of CF/PEEK instead
of cracking [4]. Therefore, it is predicted that the change in CTE for CF/PEEK will be almost like
CF/Epoxy samples.
Table 6. A partial list of CTE data obtained from the Long Duration Exposure Facility (LDEF)
experiment for CF/Epoxy samples [33].
Material Laminate Type
1
Ambient CTE 10
−6
/˚C Space CTE 10
−6
/˚C
T300/5208 Epoxy [90]
4
28.1 28.9
T300/934 Epoxy [90]
4
26.5 27.3
T300/SP-28 Epoxy [90]
4
26.3 26.8
1
The laminate is made of 4 layers with 90˚ orientation.
In the case of Ten-Koh satellite, the number of cycles can be up to 5290 for a year based on an
orbital period of approximately 98 min. For comparison with previous studies, 100 thermal cycles are
equivalent to approximately 7 days of mission duration of Ten-Koh. Table 7 list down the number of
cycles for different mission duration of Ten-Koh. The variation in CTE values with temperature
remains similar between the first month and after four months as shown in Figure 14. These findings
supported previous ground tests and in-orbit data showing that the CTE value for CF/PEEK remains
invariant up to one year.
Table 7. The number of thermal cycles for different mission durations of Ten-Koh.
Mission Duration (months) Number of Thermal Cycles
4 1763
8 3527
12 5290
There have been material science experiments performed onboard small satellites. To our
knowledge, MM is the first material science experiment that studies the effect of the space
environment on the dimensional stability of composites using small satellites. Moreover, MM
performs in situ measurements and transmits data to the ground station. The NASA LDEF studied
the effect of the space environment on the dimensional stability of composites. However, the
experiment could not transmit data in real-time nor was it performed on a small satellite [35].
The MM previous results showed that in situ observation in small form factor coupled with real-
time data transmission for a material science experiment is feasible using a small satellite platform.
Moreover, the use of commercial off the shelf components (COTS) for MM provides a viable low-cost
option for researches interested in performing in-orbit material science experiments.
Furthermore, the strain gauge attachment method has possible applications for structural health
monitoring of space structures. The fusion welding or thermal welding of strain gauge provides a
Aerospace 2020, 7, 35 20 of 24
solution to the limitations of adhesive outgassing in a standard strain gauge attachment method. The
primary influence of thermal cycling is to induce microcracking. The thermal cycling can be
considered as a low amplitude thermal fatigue resulting in microcracking changing with time [5].
The measurement methodology used in MM can be applied to monitor possible evidence of
microdamage in space structures.
4.1. Issues
4.1.1. Loss of Communication with Ten-Koh
There was no additional in-orbit data after 4.5 months due to a loss of communication between
Ten-Koh and the KIT ground station on 19 March 2019. The last data from the MM was received on
18 March 2019 between 15:15:43 and 15:16:24 Coordinated Universal Time (UTC). The last decoded
data indicated that the experiment was functioning within normal parameters. An earlier
investigation revealed that the Ten-Koh failure was likely due to radiation damage triggered by a
single event effect. Ten-Koh travelled twice over the South Atlantic anomaly region before a loss of
signal on 19 March 2019. Moreover, there was significant geomagnetic activity on 17 March 2019. The
earlier investigation explained that the geomagnetic activity had likely caused a disturbance in the
trapped radiation over the South Atlantic anomaly region. This, in turn, caused a single event effect
that may have caused Ten-Koh failure [34]. On 14 May 2019, Ten-Koh briefly re-established limited
communication. However, further material mission data and other onboard experiments data were
not received. On 4 September 2019, communication was again loss between Ten-Koh and ground
station. Currently, mission operations are still performed in the possible event that Ten-Koh re-
establishes communication.
4.1.2. UV Sensor
Initially, two UV sensors were to be installed on the external PCB. The purpose of the UV sensors
was to measure UV intensity and to compare with readings from the ISS since Ten-Koh is orbiting at
a different altitude.
One sensor that can only detect UV-C will be enclosed in the aluminium box and another sensor
that can detect all UV wavelengths will be located adjacent to the box. Figure 2 showed the location
of both UV sensors. The window of the aluminium box was planned to be composed of two UV
filters. The filters function to filter out UV-A and UV-B, thus allowing the enclosed UV sensor to focus
on measuring UV-C radiation. The UV-C wavelength in LEO is between 200 nm and 280 nm with a
mean energy of 122.6 Kcal/mole or 4.4 eV. UV-C has sufficient energy to break several chemical bonds
thus causing potential sample degradation [36]. This was the reason for applying UV filters for one
of the UV sensors.
By being selected as a secondary payload, the delivery time was primarily dependent on the
primary payload. The satellite had to be flight-qualified within a constrained schedule. This limits
the development time and reduces further tests during assembly and integration. Unfortunately, a
crack was observed on the UV filter during a shock test of the whole of Ten-Koh structure. As a result,
the mechanical design for securing the UV filter was not qualified for flight. Changes to the
mechanical design were not feasible due to the fixed delivery time. Due to possible hazard posed to
other payloads in the event of a broken filter, both UV sensors were excluded from the flight model.
4.2. Future Work
Ground test can assist in validating in-orbit data and provide a better understanding of CTE
degradation in LEO. Previously discussed ground tests were limited to validating the concept of the
MM and for flight qualification of the MM components. The next step will be to conduct further
ground tests to expose the CF/PEEK samples to a different number of thermal cycles, fluence levels
of atomic oxygen, UV intensities and different sample heating and cooling rates. This will provide a
correlation between in-orbit data and ground data for a complete understanding of CTE degradation
Aerospace 2020, 7, 35 21 of 24
in LEO. Conducting ground tests at different sample heating and cooling rates can further confirm
the effect of variable heating or cooling rate on CTE rate of change. This can act as guidance to
consider the effect of heating or cooling rate on CTE for future design of ground tests and LEO
missions.
In the current MM architecture, an 8-bit microcontroller was used on the internal PCB to handle
the strain, temperature, UV intensity measurement and Global Navigation Satellite System (GNSS)
operations. The mentioned operations were the limit for the 8-bit microcontroller onboard flash
memory. Raw data for strain and temperature in hexadecimal are transmitted to the ground station
for calculation of CTE using a separate decoder. Each strain measurement is up to 8 decimal places.
The current capability is sufficient to meet the mission requirement for MM. However, the capability
to perform onboard calculation of CTE will promote better efficiency in mission operations. Future
work can explore the replacement of the current 8-bit microcontroller with a 16-bit microcontroller.
The higher performance microcontroller allows onboard CTE calculation in addition to handling
other payload operations.
5. Conclusions
The LEO environment exposes spacecraft to factors that can degrade the dimensional stability
of the structure. The advent of high-performance CF/PEEK may limit changes in dimensional
stability. However, there are limited in-orbit data on the performance of CF/PEEK. Factors
contributing to the limitations include limited access to space, limited sample retrieval options and
difficulty in simulating actual space environment for ground test. The emerging small satellite market
provides a promising material science research platform to address the mentioned limitations. In this
paper, the Ten-Koh satellite included a material science experiment termed MM. This experiment
could perform in situ measurements of CTE for CF/PEEK samples in LEO. Strain gauges and
temperature sensors were used to provide strain and temperature measurements. The data could
then be transmitted to the ground station for calculation of CTE.
A ground validation test was performed to validate the design of MM. Analysis of the ground
test data has shown the feasibility of the MM CTE measurement system. In terms of in-orbit data, the
acquired temperature data were within the measurement range of the solar panel temperature
sensors. The CTE data exhibit a non-linear temperature dependence and varies between each sample.
Strain gauge misalignment has been identified as the reason for the marginal difference between each
sample. Comparison of in-orbit data with ground test revealed minor differences in CTE value over
a range of temperatures. This was due to difference in the sample heating and cooling rate and strain
gauge misalignment with respect to the fibre direction. Analysis of CTE data over four months was
performed before a loss of communication with Ten-Koh after four and a half months. Over four
months, there was no upward or downward shift in CTE values. The acquired in-orbit data was
shown to be consistent with previous ground tests and in-orbit data from NASA LDEF mission.
The MM experiment has demonstrated the ability to fill in the gaps between available ground
test and in-orbit data regarding CF/PEEK dimensional stability performance in LEO. Moreover, MM
proved the potential of a small satellite as a platform for conducting meaningful material science
experiments.
Abbreviations
The following abbreviations were used in this manuscript:
AACS Active Attitude Control System
ADC Analogue-to-Digital Converter
AO Atomic Oxygen
CF Carbon Fibre
CFRTP Carbon Fibre Reinforced Thermoplastic
COTS Commercial Off the Shelf Components
Aerospace 2020, 7, 35 22 of 24
CTE Coefficient of Thermal Expansion
ESA European Space Agency
FM Flight Model
GNSS Global Navigation Satellite System
ISS International Space Station
JAXA Japan Aerospace Exploration Agency
KIBO Japanese Experiment Module on ISS
KIT Kyushu Institute of Technology
LDEF Long Duration Exposure Facility
LEO Low Earth Orbit
LOC Lab-on-a-Chip
MEMS Microelectromechanical Systems
MM Material Mission
NASA National Aeronautics and Space Administration
OBC Onboard Computer
PAN Polyacrylonitrile
PCB Printed Circuit Board
PEEK Polyether Ether Ketone
PMC Polymer Matrix Composite
SP Solar panel
TMA Thermomechanical Analyzer
TVAC Thermal Vacuum Chamber
UTC Coordinated Universal Time
UV Ultraviolet
Author Contributions: Conceptualization, F.A., I.F. and N.U.; methodology, F.A., I.F. and N.U.; validation, F.A.,
I.F. and N.U.; formal analysis, F.A. and N.U.; investigation, F.A. and N.U.; resources, F.A., I.F., N.U. and K.O..;
data curation, F.A. and N.U.; writing—original draft preparation, F.A.; writing—review and editing, F.A. and
K.O.; visualization, F.A.; supervision, K.O.; project administration, K.O.; and funding acquisition, K.O. All
authors have read and agreed to the published version of the manuscript.
Funding: This research was partially funded by the Oita Prefectural Government, Oita Prefectural Organization
for Industry Creation and the working group for Ten-Koh development “Oita Challenger”.
Acknowledgements: This study cannot be completed without the effort and co-operation from the Ten-Koh
team.
Conflicts of Interest: 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.
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