Conference PaperPDF Available

In-Situ Monitoring of Carbon Fiber/Polyether Ether Ketone (CF/PEEK) Composite Thermal Expansion in Low Earth Orbit

Authors:

Abstract and Figures

The Low Earth Orbit (LEO) environment exposes spacecraft structures to ultraviolet (UV) radiation, atomic oxygen (AO), vacuum and temperature variation. Carbon Fiber/Polyether Ether Ketone (CF/PEEK) composite can be an ideal material to counter such factors. However, there is limited data on the performance of CF/PEEK composite in a LEO environment. The lack of data can be attributed to limited opportunities to send samples to space, limited retrieval methods and inaccuracy in ground tests to simulate simultaneous space environment factors. This paper discusses the design of a space exposure experiment which allows in-situ measurements of changes in the coefficient of thermal expansion (CTE) of CF/PEEK composite samples in LEO. This paper will also show the ground tests carried out before launch. The experiment eliminates the need for sample retrieval by transmitting results to the ground. The experiment is a payload onboard Ten-Koh satellite. CF/PEEK samples with attached strain gauges and temperature sensors were used to measure changes in the CTE. The CF/PEEK samples were externally attached to the top part of the satellite structure. Preliminary results indicate that the space exposure experiment can measure changes in strain and temperature required to calculate CTE values.
Content may be subject to copyright.
1
In-Situ Monitoring of Carbon Fiber/Polyether Ether Ketone (CF/PEEK)
Composite Thermal Expansion in Low Earth Orbit
By Farhan ABDULLAH,1) Kei-ichi OKUYAMA,1) Isai Fajardo TAPIA1) and NaoyaURAKAMI1)
1)Department of Applied Science for Integrated System Engineering, Kyushu Institute of Technology, Kitakyushu, Japan
(Received June 21st, 2019)
The Low Earth Orbit (LEO) environment exposes spacecraft structures to ultraviolet (UV) radiation, atomic oxygen (AO),
vacuum and temperature variation. Carbon Fiber/Polyether Ether Ketone (CF/PEEK) composite can be an ideal material to
counter such factors. However, there is limited data on the performance of CF/PEEK composite in a LEO environment. The
lack of data can be attributed to limited opportunities to send samples to space, limited retrieval methods and inaccuracy in
ground tests to simulate simultaneous space environment factors. This paper discusses the design of a space exposure
experiment which allows in-situ measurements of changes in the coefficient of thermal expansion (CTE) of CF/PEEK
composite samples in LEO. This paper will also show the ground tests carried out before launch. The experiment eliminates
the need for sample retrieval by transmitting results to the ground. The experiment is a payload onboard Ten-Koh satellite.
CF/PEEK samples with attached strain gauges and temperature sensors were used to measure changes in the CTE. The
CF/PEEK samples were externally attached to the top part of the satellite structure. Preliminary results indicate that the space
exposure experiment can measure changes in strain and temperature required to calculate CTE values.
Key Words: Atomic Oxygen, Coefficient of Thermal Expansion, Carbon Fiber/PEEK, Low Earth Orbit, Ultraviolet
Nomenclature
E
: input voltage, V
e
: differential voltage, V
K
: gauge factor
R
: gauge resistance, Ω
T
: temperature, °C
α
: coefficient of thermal expansion, /°C
R
: resistance change due to strain
ε
: strain, ppm
1. Introduction
The Low Earth Orbit (LEO) environment exposes spacecraft
structures to ultraviolet (UV) radiation, atomic oxygen (AO),
vacuum and temperature variation. These factors can affect the
dimensional stability of a spacecrafts structure. In the case of
polymer matrix composites (PMC), main factors affecting the
dimensional stability are moisture resistance, thermal
expansion, mechanical loading, and micro yielding.1) Thermal
expansion is caused by repeated thermal cycling due to the
orbital movement around the Earth. Micro yielding is caused
by microcracking in the PMC. Repeated thermal cycling can
induce microcracking.2)
Application of high-performance PMC in spacecraft
structures is crucial to limit changes to dimensional stability.
Carbon Fiber/Polyether Ether Ketone (CF/PEEK) composite is
an ideal material due to its inherent dimensional stable
properties in addition to high radiation resistance.2) PEEK is a
semi-crystalline thermoplastic polymer, and the structure is
illustrated in Fig. 1. The toughness of PEEK provides excellent
resistance to thermal cycling. 2)
Fig. 1. Chemical structure of PEEK.4)
In the Okuyama laboratory in the Kyushu Institute of
Technology, CF/PEEK was used as the primary material for the
external structure of Shinen-2 deep space probe.3) The structure
survived the launch and space environment. However, more
space data is required to further understand the behavior of
CF/PEEK including dimensional stability in space. There is
limited data on the performance of CF/PEEK composite when
exposed to the LEO environment.5) Previous sources of
CF/PEEK data primarily originated from ground tests. The lack
of space performance data can be attributed to limited
opportunities to send samples to space, limited retrieval
methods and inaccuracy in ground tests to simulate
simultaneous space environment factors.5)
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 also discusses the design
of a space exposure experiment which allows in-situ
measurements of changes in the coefficient of thermal
expansion (CTE) of CF/PEEK composites samples in LEO.
The space exposure experiment is termed as a material mission
under the Ten-Koh satellite project. 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
2
ground station. The experiment is a payload onboard the Ten-
Koh satellite which was successfully launched on 29th October
2018. This paper will also show the ground test and preliminary
space data from material mission.
2. Overview of Ten-Koh Satellite
The Ten-Koh satellite was developed by the Okuyama
Laboratory which is part of the Department of Applied Science
for Integrated System Engineering, Kyushu Institute of
Technology. Ten-Koh was launched on 29th October 2018
using a H-IIA rocket from Tanegashima Space Center. Ten-
Koh is a quasi-spherical satellite with an external structure
made primarily of Carbon Fiber Reinforced Plastic (CFRP), as
illustrated in Fig. 2. The satellite has a total mass of 22.0 kg.
The dimension of the satellite is 465 mm × 500 mm × 500 mm.
Ten-Koh orbits the Earth in a sun-synchronous sub-recurrent
orbit at an altitude of approximately 600 km.
Fig. 2. Ten-Koh flight model (FM).
The primary mission objectives for Ten-Koh are listed below:
i. To characterize the plasma environment around a
spinning spacecraft.
ii. To detect MeV-range electrons in Low Earth Orbit
(LEO):
a. Detection of MeV- range electrons under solar
storm and non-solar storm conditions.
b. Estimation of satellite anomalies when using
lightweight structures.
iii. To investigate the change of physical properties of
Carbon Fiber Reinforced Thermoplastics (CFRTP)
material samples exposed to outer space conditions.
The third primary mission objective is the focus of the
material mission.
3. Design of Material Mission
3.1. Design overview
The material mission consisted of two main components; the
internal Printed Circuit Board (PCB) and the external PCB.
Both PCB were located in the upper section of the Ten-Koh
satellite. The external PCB was installed on the top plate of the
external satellite structure for maximum field of view as
illustrated in Fig. 3. The external PCB contains the CF/PEEK
samples, strain gauges and temperature sensors. The box on the
external PCB was meant to house a UV sensor and two UV
filters. Unfortunately, the UV sensor circuit had technical
problem thus was omitted from the FM design. The internal
PCB contains the necessary components for operating the
material mission. Figures 4 and 5 illustrate the internal and
external PCB respectively.
Fig. 3. Location of material mission external PCB.
Fig. 4. Material mission internal PCB.
Fig. 5. Material mission external PCB.
3
3.2. Mechanical design
Figure 6 illustrates a detailed computer-aided design (CAD)
model of the external PCB. The external PCB measured 78 mm
long and 80mm wide. Three CF/PEEK samples were bolted to
the external PCB. Each sample consisted of two pieces of
CF/PEEK thermally welded together. The overall dimension of
each sample was 50 mm long, 10 mm wide and 2 mm thick, as
illustrated in Fig. 7. The CF used was polyacrylonitrile (PAN)
based plain woven carbon fabric manufactured by Toray. The
PEEK resin was manufactured by Victrex. In a previous
satellite developed by Okuyama laboratory termed Shinen-2,
plain woven CF/PEEK was used as the main external structure
material.3) The main reason was to maintain a quasi-isotropic
property for the external structure. The previous mission had
no in-situ measurement technique to observe the behavior of
CF/PEEK in a space environment. One sample was coated with
AO protective coating while another sample was coated with
UV protective coating. The AO protective coating used was
silsesquioxane. The UV protective coating used was a white
substance termed as yttria or yttrium oxide. The third sample
had no protective coating. The purpose of the protective coating
was to observe possible differences in the changing trend of
CTE if one of the space environment factors is excluded.
Fig. 6. CAD model of material mission external PCB.
Fig. 7. CF/PEEK sample dimension.
Each sample contains a single 0°/90° 2-element rosette
stacked type strain gauge manufactured by Tokyo Sokki
Kenkyujo. Each strain gauge was thermally welded between
the two pieces of CF/PEEK in each sample using a hot-press
machine. The completed sample is illustrated in Fig. 8. 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.
Fig. 8. CF/PEEK sample with strain gauge.
An elevated temperature test was performed to determine the
viability of the strain gauge attachment method. Strain
measurements from two identical CF/PEEK samples with two
different attachment methods were compared using a high
temperature tensile testing machine. During the test,
temperature varied between 16 °C and 106 °C to simulate the
expansion and contraction of the CF/PEEK sample. Strain
measurement showed that the trend in strain variation from the
thermal welded method was consistent with the adhesive
method. The measurement difference was approximately 5%
between both methods. Therefore, thermal welded method was
a viable attachment method. Figure 9 illustrates the result from
the strain gauge attachment method test.
Fig. 9. Difference in strain between two attachment
methods.
Temperature measurement for each sample was measured
using an integrated circuit temperature transducer or
temperature sensor attached underneath the external PCB. For
the material mission, one AD590 temperature sensor was
attached underneath each sample. The AD590 sensor is capable
of recording temperatures between -55°C to 150°C. In the
vacuum environment of space, heat was conducted via thermal
grease and lambda gel to the temperature sensor. A temperature
test was performed in a thermal vacuum chamber to compare
temperature measurement between thermocouples and the
temperature sensors. Figure 10 illustrates the temperature test
results. Test results showed an increase in temperature reading
4
from the sensors when the ambient temperature in the chamber
was increased. The gap between the thermocouple and
temperature sensor measurements was approximately 3.5 °C.
Figure 11 illustrates the cross-section sensor assembly for each
sample.
Fig. 10. Difference in temperature data between temperature
and thermocouple.
Fig. 11. Cross section sensor assembly for each CF/PEEK
sample.
3.3. Electrical design
The material mission architecture consisted of an external
PCB and an internal PCB, as illustrated in Fig. 4 and 5
respectively. The external PCB consisted of three CF/PEEK
samples. Each sample has one strain gauge and a temperature
sensor attached to it, as illustrated in Fig. 11. The main
components in the internal PCB were the Wheat stone bridge
circuit for the strain gauge, analog to digital converter (ADC)
for strain gauges and temperature sensors, secure digital (SD)
card and a PIC16F877 microcontroller for handling commands.
The strain gauge measures strain due to variation in electrical
resistance in the resistance wire.6) The variation is measured
based on Eq. (1).


(1)
Small changes in resistance are measured using a Wheat stone
bridge circuit. The bridge circuit outputs a corresponding
voltage for the change in resistance. Using Eq. (2), the strain
can then be calculated.

 (2)
A differential amplifier was used to amplify the differential
voltage. A 24-bit ADC was used to provide digital value. Each
2-element strain gauge was connected to two ADC with one
ADC for each gauge element. The strain gauge circuit block
diagram is illustrated in Fig. 12.
Fig. 12. Strain Gauge Circuit Block Diagram.
The temperature sensor used was AD590. The temperature
measurement range is from -55°C to 150°C. Three temperature
sensors were connected to one 12-bit ADC. The calculation of
CTE is based on Eq. (3).

 (3)
The reference temperature was set to room temperature. The
general process flow for material mission starts when the
ground station transmit command to internal PCB to read
temperature and strain measurements via onboard computer
(OBC). Data is then transmitted to the ground station. A
decoder program developed using C language convert the data
packets into strain, temperature and CTE values. Figure 13
illustrates the overall material mission process flow. There are
two operation modes which are normal mode and real time
mode. In normal mode, the material mission circuit will
perform three measurements with a one-minute interval. After
each measurement, data will be stored in the onboard SD card
for later downlink operations. In real time mode, the user can
set the number of measurements. The data will then be
downloaded in real time instead of being stored in the SD card.
Fig. 13. Material mission process flow.
Originally, there were two UV sensors, one sensor can detect
all UV wavelengths and the other one can detect only UV-C.
The UV filters were meant to filter out UV-A and UV-B, thus
allowing the UV sensor to focus on measuring the amount of
UV-C radiation. UV-C has the highest energy to caused
significant degradation, thus the reason for applying UV filters
in one of the sensors. However, the UV sensors were excluded
from external PCB due to technical problems and limited
development time.
0
5
10
15
20
25
30
35
40
45
0.0 500.0 1000.0 1500.0 2000.0 25 00.0 3000.0 3500.0 4000.0 4500. 0
Temperature (°C)
Time (s)
Average Temperature (Temperature Sensor)
Thermocouple
5
4. Ground Test
A ground test using a small vacuum chamber was carried out
to determine the performance of the completed material
mission circuit in a simulated space environment.
4.1. Test facility
The test was carried out in a thermal vacuum chamber
(TVAC) at the Kyuhsu Institute of Technology Center for
Nanosatellite Testing. The TVAC can produce vacuum
conditions up to 1.0x10-5 Pa.
4.2. Test procedure
Identical external and internal PCB designed for FM model
were used for the ground test. Thermocouples were attached to
the sample and temperature sensors for comparison purposes.
The test setup is illustrated in Fig. 14. The pressure in the
TVAC was set to below 1.3x10-3 Pa. The range of temperature
change was between -40°C to 22°C. This temperature range
was selected based on recorded temperatures 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°C
to -40°C at a cooling rate of -1°C/min. In the second scenario,
the test chamber temperature was increased from -40°C to 22°C
at a rate of 1°C/min.
Fig. 14. Material mission ground test setup.
5. Results and Discussion
5.1. Ground test
The results from the ground test are illustrated in Fig. 15. The
CTE values were compared with reference CTE data obtained
using thermomechanical analysis (TMA). Sample 1 refers to
CF/PEEK with no protective coating, sample 2 refers to
CF/PEEK with UV protective coating and sample 3 refers to
CF/PEEK with AO protective coating. Sample 3 provides the
closest value to the reference data with a difference in value
ranging from 0.7 to 2.4 ppm/K. The difference was because
samples used for TMA were not thermally welded. The thermal
welding process may have distorted the fiber direction. Sample
1 and 2 have different CTE values compared with sample 3. A
possible reason is that the strain gauge alignment in samples 1
and 2 is not closely aligned with the fiber direction. Even
though there is a minor difference in CTE values, the pattern in
the change of CTE is consistent between material mission
circuit and reference data. Therefore, the circuit can be used to
provide in-situ measurement of CTE for CF/PEEK.
Fig. 15. CTE values obtained using TMA and material
mission circuit.
5.2. Space data
During initial operations, it was discovered that the noise in
the downloaded signal can affect the output data. Only by
increasing the number of measurements per session can the
effect of noise be decreased. Normal mode was programmed to
obtain three measurements only during each session thus it is
unable to decrease the effect of noise. Only real time mode was
executed. However, the limited satellite pass time,
communication bandwidth and operations of other onboard
experiments placed a constraint on material mission execution.
Therefore, the average number of material mission executed is
approximately one operation every two days.
Figure 16 illustrates the variation of temperature against time.
The temperature varied approximately between -40°C to 20°C.
This temperature range was used for the ground test discussed
in the previous section. In addition, all samples displayed
approximately similar temperature readings. In the first month
of the mission, the temperature showed readings approaching -
45°C. The temperature variation was larger in the first month
compared to later months. A possible explanation was due to
increase in the speed of rotations of Ten-Koh during initial orbit
phase before stabilizing after the first month. The temperature
measurements were different because the missions were
executed during different times.
The strain measurements with respect to temperature showed
a similar change pattern with the test in the strain gauge
attachment method illustrated in Fig. 9. The strain decreased
with increasing temperature, as illustrated in Fig. 17. The CTE
for sample with no coating and UV coating varied between 0
and 8 ppm/°C, as illustrated in Fig. 18. There is an average
difference of approximately 0.8 ppm/°C between sample with
no coating and UV coating. Potential reason for the difference
External PCB
6
was due to minor strain gauge misalignment with the fiber
direction in sample with UV coating. The CTE measurements
were also compared with ground test data, as illustrated in Fig.
19. Temperatures below 0°C tend to show a similar CTE trend
between ground and space data but with a difference between
1 to 2 ppm/°C. Above 0°C, the space data showed an increasing
trend in CTE when compared with ground data. In terms of
CTE changes, Fig. 20 illustrates that the variation of CTE
measurements with temperature remain unchanged up to 120
days or 4 months. This implies that CF/PEEK is dimensionally
stable in the short-term period.
Fig. 16. Variation of sample temperature.
Fig. 17. Variation in strain with temperature.
Fig. 18. CTE measurement for all samples.
Fig. 19. Comparison of CTE measurement between ground
test data and space data.
7
Fig. 20. Comparison of CTE measurement for different
mission duration.
5. Conclusion
The LEO environment exposes spacecraft structures to UV
radiation, AO, vacuum and temperature variation. These
factors can affect the dimensional stability of spacecraft
structures. CF/PEEK composite has inherently dimensional
stable properties. However, there is limited data on the space
performance of CF/PEEK composite in LEO environment. The
lack of data can be attributed to limited opportunities to send
samples to space, limited retrieval methods and inaccuracy in
the ground test to simulate simultaneous space environment
factors. The material mission experiment allows in-situ
measurements of changes in the CTE of CF/PEEK composite
samples in LEO. The experiment utilizes strain gauges and
temperature sensors to provide data for CTE calculation.
Various ground tests have been carried to validate the design of
the experiment. There are a few design items that may affect
the accuracy of CTE measurements. In terms of space data,
initial results have shown that the variation of CTE values with
temperature remains the same up to 120 days. Overall, based
on the results of the ground tests and initial space data, it can
be concluded that the material mission experiment can be used
to measure changes in CTE for CF/PEEK in a LEO
environment.
Acknowledgments
This research was supported by the Oita Prefectural
government, Oita Prefectural Organization for Industry
Creation and the working group for Ten-Koh development
“Oita Challenger”. This study cannot be completed without the
effort and co-operation from the Ten-Koh team.
References
1)
Bowles, D. E. and Tenney, D. R.: Thermal Expansion of
Composites: Methods and Results, Large Space Systems
Technology 1980, Hampton, United States, 1980.
2)
Cogswell, F.N.: Thermoplastic Aromatic Polymer Composites,
Butterworth-Heinemann, Oxford, 1992.
3)
Fujii, H., Mashima, Y., Hibino, S., Inoue, I. and Okuyama, K.: A
New Lightweight Structure for a Nano Deep Space Probe, 16th
European Conference on Composite Materials, Seville, Spain,
ECCM16, 2014.
4)
Nakamura, T., Nakamura, H., Fujita, O., Noguchi, T. and Imagawa,
K.: The Space Exposure Experiment of PEEK sheets under Tensile
Stress, JSME International Journal, 47 (2004), pp. 365-370.
5)
Tagawa, M. and Yokota, K.: Issues and Consequences of Space
Environmental Effect on Materials, Transactions JSASS Space
Technology Japan, 7, ists26 (2009), pp. Tr_2_21Tr_2_26.
6)
Tokyo Measuring Instruments Lab: Strain Gauges, Tokyo, 2017, pp.
79.
... However, only a few polymers have been used in fiber-reinforced form. For example, in thermoset carbon epoxy and bismaleimide [14][15][16][17][18] and in thermoplastic matrices, only PEEK has been used [7,[19][20][21][22][23][24]. The other thermoplastic polymers have not been used yet. ...
... However, this can be minimized by an optimized selection of fiber and polymer matrix. For example, choosing carbon fiber and PEEK polymer creates the overall CTE of the composite in the range of 0 to 8 × 10 −6 /°C [20]. The overall CTE of graphite epoxy can also be altered from −0.06 to −2 × 10 −6 /°C by changing the fiber layup [24]. ...
... However, this can be minimized by an optimized selection of fiber and polymer matrix. For example, choosing carbon fiber and PEEK polymer creates the overall CTE of the composite in the range of 0 to 8 × 10 −6 / • C [20]. The overall CTE of graphite epoxy can also be altered from −0.06 to −2 × 10 −6 / • C by changing the fiber layup [24]. ...
Article
Full-text available
This review paper discusses the effect of polymers, especially thermoplastics, in environments with low earth orbits. Space weather in terms of low earth orbits has been characterized into seven main elements, namely microgravity, residual atmosphere, high vacuum, atomic oxygen, ultraviolet and ionization radiation, solar radiation, and space debris. Each element is discussed extensively. Its effect on polymers and composite materials has also been studied. Quantification of these effects can be evaluated by understanding the mechanisms of material degradation caused by each environmental factor along with its synergetic effect. Hence, the design elements to mitigate the material degradation can be identified. Finally, a cause-and-effect diagram (Ishikawa diagram) is designed to characterize the important design elements required to investigate while choosing a material for a satellite’s structure. This will help the designers to develop experimental methodologies to test the composite material for its suitability against the space environment. Some available testing facilities will be discussed. Some potential polymers will also be suggested for further evaluation.
Article
The effects of electron beam irradiation were examined on perforation holes and fragments from quasi-isotropic carbon fiber reinforced plastic plates (CFRP plates). CFRP plates with a total dose up to 60 MGy of electron beam irradiation were prepared for hypervelocity impact experiments. The perforation holes of CFRP plates in the hypervelocity impact experiments were decreased with electron beam irradiation when aluminum alloy spherical projectiles (2017-T4) of diameter of 1 mm struck at an impact velocity of 5 km/s. However, a slight reduction in the size of the perforation hole was observed at 2 km/s. Fragments ejected from CFRP plates were examined after collection from the impact side and rear side of the targets. The cumulative number distributions of fragment length and aspect ratio were compared among specimens with different total doses.
Article
The importance of space environmental effect on material is addressed. Some examples of the material degradation in a particular space environment are introduced. In order to endorse the material properties requested in a mission, ground-based studies are quite important. However, present deficient ground-based simulation technology cannot perform the accurate assessment of the material degradation in space. On the other hand, infrastructures for the material evaluation using Kibo and International Space Station, is now established after long desire of material scientists and engineers. However, we will face difficulty for retrieve the exposed sample by the retirement of Space Shuttle. Development of new material evaluation methods using small satellite without retrieving the samples and of the accurate ground-based simulation techniques are both highly desired to achieve low-risk missions.
Article
To find out the degradation behavior of polymer in the real space, space exposure experiments utilizing the International Space Station (ISS) were scheduled. PEEK sheets under tensile stresses were exposed to the environment around the ISS orbit, and were irradiated by atomic oxygen (AO), ultraviolet ray, and electron beam (EB) in the ground test facility., This study introduces the outline of these experiments, and shows the results of AO and EB pilot irradiation tests as follows: (1) Test piece surfaces after AO exposure exhibited significant morphological damages characterized by micron-sized conical pits. (2) Thickness reductions of the test pieces by AO exposure increased with increasing tensile stress. (3) Residual strength after AO exposure could be estimated by taking account of thickness reduction. (4) No significant change was observed on surface morph, mass, chemical structure, and tensile properties of the test pieces after EB exposure regardless of tensile stress.
Thermal Expansion of Composites: Methods and Results, Large Space Systems Technology -1980
  • D E Bowles
  • D R Tenney
Bowles, D. E. and Tenney, D. R.: Thermal Expansion of Composites: Methods and Results, Large Space Systems Technology -1980, Hampton, United States, 1980.