- Email: [email protected]
Performance of EDT system for deorbit devices using new materials
Acta Astronautica xxx (xxxx) xxx–xxx
Contents lists available at ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Performance of EDT system for deorbit devices using new materials Tsuyoshi Satoa,∗, Satomi Kawamotob, Yasushi Ohkawab, Takeo Watanabea, Koh Kamachic, Hiroshi Okuboa a b c
Department of Mechanical Engineering, Kanagawa Institute of Technology, 1030, Shimo-ogino, Atsugi, Kanagawa, 243-0292, Japan Research and Development Directorate, Japan Aerospace Exploration Agency, 7-44-1 Jindaiji-higashi-machi, Chofu, Tokyo, 182-8522, Japan Research and Development, ALE Co., Ltd., 2F, 2-11-8 Shiba-Daimon, Minato-ku, Tokyo, 105-0012, Japan
ARTICLE INFO
ABSTRACT
Keywords: Electrodynamic tether Post mission disposal Microsatellite Orbital debris
A deorbit device is required for some microsatellites to meet space debris mitigation guidelines, although very challenging in terms of limited resources and reliability. Many groups are conducting research on post-mission disposal (PMD) devices using an electrodynamic tether (EDT) due to its high efficiency and simplicity. Since an EDT for microsatellites must be lightweight, with some strength, high conductivity, high survivability, and meet other requirements, such new materials as carbon nanotube yarn, metal-plated fiber, and metal-deposited thin film are assumed for a tape type tether. In order to determine the appropriate EDT dimensions such as tether width and length, the deorbit capabilities must be evaluated by numerical simulations in advance, as the thrust obtained varies depending on the EDT dimensions, orbital parameters, and other factors. The required resources of the EDT system such as mass and electric power can then be obtained for each orbit, satellite, and deorbit time. Thus, several prototypes of tape type tethers were made and evaluated in various tests.
1. Introduction There has been active space development in recent years, with a growing number of satellites being launched. Many microsatellite missions are also being planned or prepared. Accordingly, measures against space debris are urgently needed, and countries are working on measures against space debris from various viewpoints such as deorbit, protection, and observation [1–3]. The space debris mitigation guidelines of the Inter-Agency Space Debris Coordination Committee (IADC) recommend the deorbiting of microsatellites launched into LEO within 25 years after their mission ends [4]. A deorbit device is required at high altitude where orbital lifetime is long enough to meet the space debris mitigation guidelines, although very challenging in terms of limited resources, reliability, and longterm storage in space. Many groups are conducting research on a post-mission disposal (PMD) device using an electrodynamic tether (EDT) [5–9]. Fig. 1 shows the concept of the PMD device. First, after the mission, the folded tape type tether is deployed by releasing an end mass. After deployment, the tape type tether crosses Earth's magnetic field where induced electromotive force is generated, and then a potential differ-
∗
ence develops in the tether. Electrons are collected by the bare surface of the tape type tether in the region where the potential is positive with respect to ambient plasma, whereas ions are collected and electrons are emitted from the other end of the tether where the potential is negative (Fig. 2). As a result, electromagnetic force is generated by the current flowing in the tether and Earth's magnetic field, and the satellite starts to decelerate. In addition, deceleration due to atmospheric drag can be fully expected. An EDT has two major advantages. The first advantage is that the EDT derives power from the geomagnetic field, and thus needs no propellant. The second advantage is that EDT does not require thrust vector control, and there are no restrictions on the installation location [10,11]. In this study, we focus on the EDT to be used as a PMD device of microsatellites. It is necessary to determine appropriate tether specifications for such requirements as satellite weight and orbital conditions. The materials and configuration of the tether are investigated, and then the effects of tether parameters are evaluated by numerical simulations. Various effects such as dynamics of tether are considered in this simulation. Then, tape type tethers are made of several new materials based on the simulation results, and various tests such as tensile test are conducted to evaluate those tethers.
Corresponding author. E-mail address: [email protected] (T. Sato).
https://doi.org/10.1016/j.actaastro.2020.01.029 Received 30 July 2019; Received in revised form 20 December 2019; Accepted 20 January 2020 0094-5765/ © 2020 IAA. Published by Elsevier Ltd. All rights reserved.
Please cite this article as: Tsuyoshi Sato, et al., Acta Astronautica, https://doi.org/10.1016/j.actaastro.2020.01.029
Acta Astronautica xxx (xxxx) xxx–xxx
T. Sato, et al.
folded state and deployed on orbit. A folded tape type tether can be easily deployed by ejecting the end mass where the end of the tether is attached. Regarding the deployment of the folded tape type tether, it has been successfully deployed in 2010 in a sounding rocket experiment [14]. And since the releasing mechanism can be miniaturized, such a tether is suitable for microsatellites. In this concept, flexibility and ease of folding are important as the tether is folded and stored. In the previous study, aluminum and polyester were used as the materials of a flat tape type tether (called “ALPET”) that featured a
Nomenclature
J (x ) (x ) e np H me
Current Potential Elementary electron charge Plasma number density Tether width Electron mass
Fig. 1. Concept of PMD device using EDT.
sandwich structure (Fig. 3). This tether has sufficient conductivity and strength by using polyester as the strength material, and aluminum as the conductive material. The ALPET tether has already been successfully deployed in an experiment in space (T-Rex, 2010) [14]. The tether was installed on a sounding rocket and achieved successful deployment of a 136-m tether in space. However, the ALPET tether is rather heavy. Therefore, in this research, we focused on new materials such as metaldeposited thin film, carbon nanotube yarn, and silver-plated nylon yarn (ODEX). The use of these materials is expected to reduce the tether weight. Tape type tethers are roughly divided into film-like tethers and ribbon-like tethers. The film-like tether can be made thinner, with a reduced volume when folded. The ribbon-like tether is made by knitting fine fiber material into a mesh or weaving fine fiber material into an organdy. Thus, it is difficult for cracks to develop in a ribbon-like tether, and high survivability is expected. Moreover, making the mesh spacing coarser can be expected to reduce the tether weight. The thrust obtained varies depending on the tether width, tether length, and other factors. In order to determine the dimensions of the tether, the deorbit capabilities must be evaluated by numerical simulations in advance.
2. Overview of EDT system for microsatellites 2.1. Tape type tether Generally, an EDT for microsatellites must be lightweight, with some strength, high conductivity, high survivability, and meet other requirements. Given their high electron collection performance, tape type tethers are suitable for microsatellites even if the tethers are short in length. Tape type tethers are also expected to survive debris collisions [12,13]. It is assumed that the tape type tether is installed in a
2.2. Electron emitter Typically, on electron emitter is installed on the end of the tether where the potential is negative. When the electron emitter is attached, more electrons are emitted and the potential distribution of the tether moves to the positive side. As a result, more electrons can be collected to the tether surface, thereby increasing the current flowing through the tether. The presence or absence of an electron emitter greatly affects EDT performance. Using an electron emitter can shorten the time required for deorbiting, and thus reduce the risk of debris collision during PMD device operation. Moreover, a collision avoidance maneuver may be possible by turn on/off the electron emitter. In case of unexpected
Fig. 2. Model of electron emission and collection. 2
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Fig. 3. Schematic drawing of the ALPET tether.
failure, deorbit can still be performed given the slight flow of current enabled by passive ion collection and electron emission from the bare tether. The field emission cathodes (FEC) using carbon nanotube as an electron emitter have been studied by JAXA [15]. The FEC is considered optimal for EDT systems because it has a very simple system that only needs to be energized. Furthermore, high mechanical strength, a high aspect ratio, and chemical stability can be expected by using carbon nanotube as the emitter material. Fig. 4 shows a schematic drawing of the carbon-nanotube-based FEC. The effectiveness of the FEC in low Earth orbit (LEO) has been confirmed by the Kounotori Integrated Tether Experiment (KITE)—an on-orbit demonstration experiment conducted by JAXA in 2017 using HTV-6 [16]. The results of KITE showed that carbon-nanotube-based FEC can be expected to operate for several years by protecting it from direct exposure to atomic oxygen (AO) [17]. And efforts are now being made to improve FEC performance for use in microsatellites. Fig. 5 shows an example of the laboratory model FEC, which performs 80-mA-level electron emission from a 90 × 90 mm2 area. The required weight and area of the electron emitter are determined by the maximum emission current, and thus, the effect of maximum emission current on performance must be investigated by numerical simulations.
conductivity is assumed to be the theoretical value of aluminum. The satellite was assumed to weigh 60 kg and measure 80 cm × 60 cm × 60 cm in size. 3.1. Effects of orbit The effects of orbit on performance were evaluated without an electron emitter. Figs. 9 and 10 show the electromagnetic force and atmospheric drag for each orbital altitude at an orbital inclination of 98.4°. The initial altitude of the satellite was 800 km. Fig. 9 shows in case of the tether width is 150 mm, the tether length is 500 m, and Fig. 10 shows in case of the tether width is 25 mm, the tether length is 900 m. Fig. 11 shows in case of the tether width is 150 mm, the tether length is 500 m, and Fig. 12 shows in case of the tether width is 25 mm, the tether length is 900 m, both for an inclination of 25°. The parameters were chosen so that the satellite could be deorbited in a few years. For the orbital inclination angle, 25° was selected as a case where electromagnetic force is highly effective, and 98.4° as a less effective case. At high inclination, very little electromagnetic force is produced as expected, when an electron emitter isn't used. As for the tether width, 150 mm and 25 mm were selected to compare the case where the influence of atmospheric drag is large and the case where it is relatively small. As shown in Figs. 9–12, the effect of atmospheric drag clearly becomes more remarkable as the orbital altitude decreases. Although atmospheric drag is dominant at an orbital inclination angle of 98.4°, the ratio of electromagnetic force increases dramatically at an orbital inclination angle of 25°. According to the above results and previous research [20], the EDT system is less effective in a high inclination orbit where the direction of the geomagnetic field is not suitable. Hereafter, we will introduce an example where an orbital inclination angle is 98.4° as a less effective for an EDT application.
3. Evaluation of the effects on EDT performance by various parameters The effects of orbital conditions and various tether parameters on EDT system performance were investigated by numerical analysis using a simulation code developed by JAXA [18]. To consider tether flexibility, the tether is modeled as a lumped mass by dividing it into a series of point masses, each connected to its neighbors by a spring and a viscous damper (Fig. 6). The equation of motion for each point mass is formulated in a coordinate system wherein the origin is the system's center of mass in LEO (Fig. 7). Orbital perturbations caused by air drag and geopotential are considered using Gauss's vibrational equations of motion. The Naval Research Laboratory's Mass Spectrometer and Incoherent Scatter radar Extended model (NRLMSISE-00) is used for the atmosphere model, the International Reference Ionosphere (IRI) 2016 model is used for the plasma model, and International Geomagnetic Reference Field (IGRF-12) (10 × 10) is used for Earth's geopotential field. F10.7 solar flux and Ap magnetic indices as of 1999 are used, as one example. For electron collection by the bare tether, the two-dimensional orbital motion limit (OML) equation described below was used [19].
enp 2H dJ = dx
2e (x ) me
3.2. Effects on maximum emission current of electron emitter The effects on the maximum emission current of the electron emitter were evaluated. Fig. 13 shows the electromagnetic force and atmospheric drag for each maximum emission current of the electron emitter. The tether width is 25 mm and the tether length is 900 m. It shows the case at an orbital altitude of 800 km and an orbital inclination angle of 98.4°. This result suggests that, with an electron emitter, an EDT system can significantly enhance electromagnetic force.
(1)
In Eq. (1), J (x ), (x ), e, np, H , and me denote the current, potential, elementary electron charge, plasma density, tether width, and electron mass, respectively. The potential of tether is determined by iterating calculations to balance between collected and emitted electrons. The atmospheric drag of the tether depends on the projected area viewed from the orbital velocity direction. Because the attitude around the axis of the tape type tether cannot be controlled, the average projected area is used. As the thickness of the tape type tether is negligible with respect to the width, the average projected area is 2/ of the tether width. In the simulation, the tether is assumed to be aluminum with a thickness of 6 μm, and attached to one side of the ODEX ribbon (Fig. 8). The
Fig. 4. Schematic drawing of the carbon-nanotube-based FEC. 3
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3.3. Effects of tether width and length The effects of tether length and width on performance were evaluated when an emitter is not used. Fig. 14 shows the electromagnetic force and atmospheric drag for each tether width and length. It shows the case at an orbital altitude of 800 km and at an orbital inclination angle of 98.4° without an electron emitter. As shown in Fig. 14, the electromagnetic force is almost proportional to the tether width and proportional to the square of the tether length. And atmospheric drag is also proportional to the projected area. The effect of the tether length is thus considered to be greater than the effect of the tether width. Consequently, the larger the potential difference due to induced electromotive force with a long tether length, and the larger the amount of flowing current. 3.4. Effect of tether conductivity The effect of tether conductivity on electromagnetic force was evaluated. By evaluating the effect of tether conductivity, it is possible to estimate the required thickness of the tether's conductive layer. Fig. 15 shows the electromagnetic force for each tether conductivity. Here, cases of using and not using an electron emitter are assumed. An electron emitter is used with a maximum emission current of 40 mA. It shows the case at an orbital altitude of 800 km and at an orbital inclination angle of 98.4°. The tether width is 25 mm and the tether length is 900 m. As shown in Fig. 15, the electromagnetic force is saturated at a certain conductivity or more. This is because flowing current does not increase even when conductivity is good due to other conditions such as the potential difference. When an electron emitter is used, the saturation point moves to the low resistance side.
Fig. 5. Example of the laboratory model FEC.
3.5. Summary of simulation results and image of deorbit Figs. 16 and 17 show images of the changes in the orbital altitude of the satellite under several conditions relative to the effects of tether length and width, respectively with and without an electron emitter. The initial orbital altitude is 800 km and the orbit inclination angle is 98.4°. As shown in Figs. 16 and 17, it is more effective to extend the tether length than to extend the tether width, and performance can be further improved by using an electron emitter. Even if an orbital inclination angle is 98.4° as a less effective to an EDT application, the 25year rule can be sufficiently satisfied without electron emitter. Fig. 18 shows an image of the change in orbital altitude of the satellite under several inclination angles. As shown in Fig. 18, the EDT system is clearly under a challenging condition at an orbit inclination angle of 98.4°. Therefore, a somewhat elongated tether can be expected to achieve higher performance. However, an elongated tether tends to require a larger storage volume, and a balance between storage volume and performance is important. By making such evaluations, the specifications of the tether and electron emitter can be determined for such EDT requirements as orbital conditions, weight, and deorbit time.
Fig. 6. Lumped mass model.
4. Trial manufacture of tethers and various tests Several EDT prototypes were made to measure the actual tether parameters. The EDT must have specific strength, conductivity, survivability, flexibility, non-stickiness in space, and meet other requirements. Fig. 19 shows several prototype EDTs. The aluminum-deposited polyester film and aluminum-deposited polyimide film are very lightweight, but have poor conductivity due to the deposition of aluminum. Therefore, the thickness of the vapor deposition layer must be increased more than that in general vapor deposition. Given the aluminum surface of the tether, sticking in space may occur, and the surface will become hot due to the poor thermooptical properties of aluminum. Consequently, surface treatment is required.
Fig. 7. The system in LEO.
4
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Fig. 8. ODEX (silver-plated nylon yarn) ribbon.
Fig. 9. Effects of orbital altitude (tether length: 500 m tether, width: 150 mm, inclination angle: 98.4°).
ig. 10. Effects of orbital altitude (tether length: 900 m, tether width: 25 mm, inclination angle: 98.4°).
Fig. 11. Effects of orbital altitude (tether length: 500 m, tether width: 150 mm, inclination angle: 25°).
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Fig. 12. Effects of orbital altitude (tether length: 900 m, tether width: 25 mm, inclination angle: 25°).
Fig. 13. Effects on maximum emission current of electron emitter.
Fig. 14. Effects of tether width and length.
Fig. 15. Effect of tether conductivity.
The carbon nanotube ribbon is woven in a ribbon shape with carbon nanotube yarn. Carbon nanotube yarn has high specific strength and adequate conductivity. Conductivity can be increased by applying metal plating on the surface of the carbon nanotube yarn. Carbon
nanotube yarn is very flexible compared to carbon fiber. However, carbon nanotube is degraded by atomic oxygen, and thus can be expected to eventually and naturally disappear in space under long-term durations. 6
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Fig. 16. Image of the change in orbital altitude of the satellite (effect of tether length).
Fig. 17. Image of the change in orbital altitude of the satellite (effect of tether width).
Fig. 18. Image of the change in orbital altitude of the satellite (effect of orbit).
Fig. 19. Prototypes of various tethers.
The ODEX ribbon has a good specific strength and is flexible. The ODEX ribbon also has less fuzzing of fibers. But given its low conductivity, measures must be devised to improve conductivity, such as attaching an aluminum foil.
Cracks are considered difficult to develop in the ribbon-like tether, and thus high survivability is expected. The tensile test confirmed how the ribbon-like tether broke. Fig. 20 shows the appearance of the fractured surfaces of the ODEX ribbon and an aluminum-deposited 7
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Fig. 20. Appearance of fractured surfaces.
polyester film subjected to a tensile test for comparison. In the ribbonlike tether, cracks did not spread throughout the tether, and the entire tether did not break. In the future, various tests such as a tensile test and atomic oxygen irradiation test, as well as electrical resistance measurements, will be conducted.
Mitigation Guidelines, IADC Action Item number 22.4 (2007). [5] T. Watanabe, T. Sato, H. Sakamoto, S. Kawamoto, Y. Ohkawa, K. Kamachi, L. Okajima, A study on post-mission disposal device for microsatellite using electrodynamic tape tether Technology, JAXA Special Publication: Proceedings of the 8th Space Debris Workshop, 2019. [6] H. Sakamoto, K. Arai, K. Umetsu, T. Fujita, T. Sato, T. Watanabe, S. Kawamoto, K. Kamachi, Y. Oikawa, L. Okajima, Experimental study on separation of tape tether deployment module for PMD, JAXA Special Publication: Proceedings of the 8th Space Debris Workshop, 2019. [7] G. Sánchez-Arriaga, X. Chen, Modelling and performance of electrodynamic low work-function tethers with photoemission effects, J. Propul. Power 34 (1) (2018) 213–220. [8] J.K. McTernan, S.G. Bilén, Embodied energy repurposing via energy-harvesting electrodynamic tethers, J. Spacecraft Rockets 54 (4) (2017) 789–795. [9] I.C. Bell III, B.E. Gilchrist, J. McTernan, S.G. Bilén, Investigating miniaturized electrodynamic tethers for picosatellites and femtosatellites, J. Spacecraft Rockets 54 (No. 1) (2017) 55–66. [10] K. Iki, S. Kawamoto, Y. Ohkawa, T. Okumura, K. Kawashima, M. Takai, K. Izawa, K. Matsumoto, S. Suzuki, Y. Katayama, Y. Horikawa, K. Inoue, Expected on-orbit tether deployment dynamics on the KITE mission, Trans. JSASS Aerospace Tech. Japan 14 (2016) ists30 Pr_9-Pr_18. [11] M. Khadija, Results of the STEP-AIRSEDS Electrodynamic Tether Deployment Simulations, AIAA Paper 2002-1119 (2002). [12] A. Francesconi, C. Giacomuzzo, F. Branz, E.C. Lorenzini, Survavibility to hypervelocity impacts of electrodynamic tape tethers for deorbiting spacecraft in LEO, 6th European Conference on Space Debris, European Space Operations Centre, Darmstadt, Germany, April 2013, pp. 22–25. [13] K. Makihara, S. Kondo, Structural evaluation for electrodynamic tape tethers against hypervelocity space debris impacts, J. Spacecraft Rockets 55 (4) (January 2018) 1–11. [14] A.H. Fujii, T. Watanabe, H. Sahara, H. Kojima, S. Takehara, Y. Yamagiwa, et al., Space Demonstration of Bare Electrodynamics Tape-Tether Technology on the Sounding Rocket S520-25, AIAA Paper 2011-06503 (2011). [15] Y. Ohkawa, A. Izawa, Y. Yamagiwa, S. Kawamoto, S. Nishida, S. Kitamura, Research and development of carbon nanotube cathodes for electric propulsion, Trans. JSASS Aerospace Tech. Japan 8 (2010) Pb_27–Pb_32 ists27. [16] Y. Ohkawa, K. Iki, T. Okumura, S. Kawamoto, Y. Horikawa, K. Inoue, Y. Kobayashi, T. Uchiyama, T. Kasai, A quick review of an electrodynamic tether experiment on the H-II transfer vehicle, 35th Int'l Electric Propulsion Conf., Atlanta, USA, IEPC2017-359, 2017. [17] Y. Ohkawa, T. Okumura, K. Iki, H. Okamoto, S. Kawamoto, Operation of a carbon nanotube field-emission cathode in low Earth orbit, J. Vac. Sci. Technol. B 37 (2019). [18] S. Kawamoto, Y. Ohkawa, T. Okumura, K. Iki, H. Okamoto, Performance of electrodynamic tether system for debris deorbiting: Re-evaluation based on the results of KITE experiments, IAC-18- A6.6.5, 69th International Astronautical Congress (IAC), Bremen, Germany, October 2018, pp. 1–5. [19] J.R. Sanmartin, M. Martinez-Sanchez, E. Ahedo, Bare wire anodes for electrodynamic tethers, J. Propul. Power 9 (1993) Number 3. [20] T. Sato, S. Kawamoto, T. Watanabe, K. Kamachi, Y. Ohkawa, Y. Oikawa, L. Okajima, H. Okubo, A study on PMD device for microsatellites using electrodynamic tether, 2019-r-29, Proceedings of the 32nd International Symposium on Space Technology and Science, Fukui, Japan, 2019, pp. 15–21 (June).
4. Conclusion We focused on an EDT for microsatellite PMD devices, and considered the material and configuration of the tether. It was assumed that the tape type tether is installed in a folded state and deployed on orbit. By using the folded tape type tether, easy deployment could be expected by ejecting an end mass attached at the end of the tether. And since the device can be miniaturized, it is suitable for microsatellites. The effects of tether parameters were evaluated by numerical simulations. The specifications of the tether and electron emitter could be determined for such EDT requirements as orbital conditions, weight, and deorbit time. We also manufactured several tethers made of new materials based on the simulation results, and tensile tests confirmed how the tethers broke. As a result, the manner of breaking was confirmed as being different among those tethers. In the future, additional ground experiments and numerical analysis will be conducted for a more detailed study. Acknowledgements Part of this research was conducted based on the JAXA Space Innovation through Partnership and Co-creation (J-SPARC) agreement between JAXA and ALE, and the joint research agreement between the Kanagawa Institute of Technology and ALE. References [1] J.-C. Liou, The near-Earth orbital debris problem and the challenges for environment remediation, The 3rd International Space World Conference, 2012. [2] C. Colombo, A. Rossi, F.D. Vedova, A. Francesconi, C. Bombardelli, M. Trisolini, J.L. Gonzalo, P.D. Lizia, C. Giacomuzzo, S.B. Khan, R. G-Pelayo, V. Braun, B.B. Virgili, H. Krag, Effects of passive de-orbiting through drag and solar sails and electrodynamic tethers on the space debris environment, IAC-18-A6.2.8, 69th International Astronautical Congress (IAC), Bremen, Germany, 1-5 October, 2018. [3] R.P. Hoyt, I.M. Barnes, N.R. Voronka, J.T. Slostad, The terminator Tape™: a costeffective de-orbit module for end-of-life disposal of LEO satellites, AIAA SPACE 2009 Conference & Exposition, Pasadena, California, 14-17 September, 2009. [4] Inter-Agency Space Debris Coordination Committee (IADC), IADC Space Debris
8
Contents lists available at ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Performance of EDT system for deorbit devices using new materials Tsuyoshi Satoa,∗, Satomi Kawamotob, Yasushi Ohkawab, Takeo Watanabea, Koh Kamachic, Hiroshi Okuboa a b c
Department of Mechanical Engineering, Kanagawa Institute of Technology, 1030, Shimo-ogino, Atsugi, Kanagawa, 243-0292, Japan Research and Development Directorate, Japan Aerospace Exploration Agency, 7-44-1 Jindaiji-higashi-machi, Chofu, Tokyo, 182-8522, Japan Research and Development, ALE Co., Ltd., 2F, 2-11-8 Shiba-Daimon, Minato-ku, Tokyo, 105-0012, Japan
ARTICLE INFO
ABSTRACT
Keywords: Electrodynamic tether Post mission disposal Microsatellite Orbital debris
A deorbit device is required for some microsatellites to meet space debris mitigation guidelines, although very challenging in terms of limited resources and reliability. Many groups are conducting research on post-mission disposal (PMD) devices using an electrodynamic tether (EDT) due to its high efficiency and simplicity. Since an EDT for microsatellites must be lightweight, with some strength, high conductivity, high survivability, and meet other requirements, such new materials as carbon nanotube yarn, metal-plated fiber, and metal-deposited thin film are assumed for a tape type tether. In order to determine the appropriate EDT dimensions such as tether width and length, the deorbit capabilities must be evaluated by numerical simulations in advance, as the thrust obtained varies depending on the EDT dimensions, orbital parameters, and other factors. The required resources of the EDT system such as mass and electric power can then be obtained for each orbit, satellite, and deorbit time. Thus, several prototypes of tape type tethers were made and evaluated in various tests.
1. Introduction There has been active space development in recent years, with a growing number of satellites being launched. Many microsatellite missions are also being planned or prepared. Accordingly, measures against space debris are urgently needed, and countries are working on measures against space debris from various viewpoints such as deorbit, protection, and observation [1–3]. The space debris mitigation guidelines of the Inter-Agency Space Debris Coordination Committee (IADC) recommend the deorbiting of microsatellites launched into LEO within 25 years after their mission ends [4]. A deorbit device is required at high altitude where orbital lifetime is long enough to meet the space debris mitigation guidelines, although very challenging in terms of limited resources, reliability, and longterm storage in space. Many groups are conducting research on a post-mission disposal (PMD) device using an electrodynamic tether (EDT) [5–9]. Fig. 1 shows the concept of the PMD device. First, after the mission, the folded tape type tether is deployed by releasing an end mass. After deployment, the tape type tether crosses Earth's magnetic field where induced electromotive force is generated, and then a potential differ-
∗
ence develops in the tether. Electrons are collected by the bare surface of the tape type tether in the region where the potential is positive with respect to ambient plasma, whereas ions are collected and electrons are emitted from the other end of the tether where the potential is negative (Fig. 2). As a result, electromagnetic force is generated by the current flowing in the tether and Earth's magnetic field, and the satellite starts to decelerate. In addition, deceleration due to atmospheric drag can be fully expected. An EDT has two major advantages. The first advantage is that the EDT derives power from the geomagnetic field, and thus needs no propellant. The second advantage is that EDT does not require thrust vector control, and there are no restrictions on the installation location [10,11]. In this study, we focus on the EDT to be used as a PMD device of microsatellites. It is necessary to determine appropriate tether specifications for such requirements as satellite weight and orbital conditions. The materials and configuration of the tether are investigated, and then the effects of tether parameters are evaluated by numerical simulations. Various effects such as dynamics of tether are considered in this simulation. Then, tape type tethers are made of several new materials based on the simulation results, and various tests such as tensile test are conducted to evaluate those tethers.
Corresponding author. E-mail address: [email protected] (T. Sato).
https://doi.org/10.1016/j.actaastro.2020.01.029 Received 30 July 2019; Received in revised form 20 December 2019; Accepted 20 January 2020 0094-5765/ © 2020 IAA. Published by Elsevier Ltd. All rights reserved.
Please cite this article as: Tsuyoshi Sato, et al., Acta Astronautica, https://doi.org/10.1016/j.actaastro.2020.01.029
Acta Astronautica xxx (xxxx) xxx–xxx
T. Sato, et al.
folded state and deployed on orbit. A folded tape type tether can be easily deployed by ejecting the end mass where the end of the tether is attached. Regarding the deployment of the folded tape type tether, it has been successfully deployed in 2010 in a sounding rocket experiment [14]. And since the releasing mechanism can be miniaturized, such a tether is suitable for microsatellites. In this concept, flexibility and ease of folding are important as the tether is folded and stored. In the previous study, aluminum and polyester were used as the materials of a flat tape type tether (called “ALPET”) that featured a
Nomenclature
J (x ) (x ) e np H me
Current Potential Elementary electron charge Plasma number density Tether width Electron mass
Fig. 1. Concept of PMD device using EDT.
sandwich structure (Fig. 3). This tether has sufficient conductivity and strength by using polyester as the strength material, and aluminum as the conductive material. The ALPET tether has already been successfully deployed in an experiment in space (T-Rex, 2010) [14]. The tether was installed on a sounding rocket and achieved successful deployment of a 136-m tether in space. However, the ALPET tether is rather heavy. Therefore, in this research, we focused on new materials such as metaldeposited thin film, carbon nanotube yarn, and silver-plated nylon yarn (ODEX). The use of these materials is expected to reduce the tether weight. Tape type tethers are roughly divided into film-like tethers and ribbon-like tethers. The film-like tether can be made thinner, with a reduced volume when folded. The ribbon-like tether is made by knitting fine fiber material into a mesh or weaving fine fiber material into an organdy. Thus, it is difficult for cracks to develop in a ribbon-like tether, and high survivability is expected. Moreover, making the mesh spacing coarser can be expected to reduce the tether weight. The thrust obtained varies depending on the tether width, tether length, and other factors. In order to determine the dimensions of the tether, the deorbit capabilities must be evaluated by numerical simulations in advance.
2. Overview of EDT system for microsatellites 2.1. Tape type tether Generally, an EDT for microsatellites must be lightweight, with some strength, high conductivity, high survivability, and meet other requirements. Given their high electron collection performance, tape type tethers are suitable for microsatellites even if the tethers are short in length. Tape type tethers are also expected to survive debris collisions [12,13]. It is assumed that the tape type tether is installed in a
2.2. Electron emitter Typically, on electron emitter is installed on the end of the tether where the potential is negative. When the electron emitter is attached, more electrons are emitted and the potential distribution of the tether moves to the positive side. As a result, more electrons can be collected to the tether surface, thereby increasing the current flowing through the tether. The presence or absence of an electron emitter greatly affects EDT performance. Using an electron emitter can shorten the time required for deorbiting, and thus reduce the risk of debris collision during PMD device operation. Moreover, a collision avoidance maneuver may be possible by turn on/off the electron emitter. In case of unexpected
Fig. 2. Model of electron emission and collection. 2
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Fig. 3. Schematic drawing of the ALPET tether.
failure, deorbit can still be performed given the slight flow of current enabled by passive ion collection and electron emission from the bare tether. The field emission cathodes (FEC) using carbon nanotube as an electron emitter have been studied by JAXA [15]. The FEC is considered optimal for EDT systems because it has a very simple system that only needs to be energized. Furthermore, high mechanical strength, a high aspect ratio, and chemical stability can be expected by using carbon nanotube as the emitter material. Fig. 4 shows a schematic drawing of the carbon-nanotube-based FEC. The effectiveness of the FEC in low Earth orbit (LEO) has been confirmed by the Kounotori Integrated Tether Experiment (KITE)—an on-orbit demonstration experiment conducted by JAXA in 2017 using HTV-6 [16]. The results of KITE showed that carbon-nanotube-based FEC can be expected to operate for several years by protecting it from direct exposure to atomic oxygen (AO) [17]. And efforts are now being made to improve FEC performance for use in microsatellites. Fig. 5 shows an example of the laboratory model FEC, which performs 80-mA-level electron emission from a 90 × 90 mm2 area. The required weight and area of the electron emitter are determined by the maximum emission current, and thus, the effect of maximum emission current on performance must be investigated by numerical simulations.
conductivity is assumed to be the theoretical value of aluminum. The satellite was assumed to weigh 60 kg and measure 80 cm × 60 cm × 60 cm in size. 3.1. Effects of orbit The effects of orbit on performance were evaluated without an electron emitter. Figs. 9 and 10 show the electromagnetic force and atmospheric drag for each orbital altitude at an orbital inclination of 98.4°. The initial altitude of the satellite was 800 km. Fig. 9 shows in case of the tether width is 150 mm, the tether length is 500 m, and Fig. 10 shows in case of the tether width is 25 mm, the tether length is 900 m. Fig. 11 shows in case of the tether width is 150 mm, the tether length is 500 m, and Fig. 12 shows in case of the tether width is 25 mm, the tether length is 900 m, both for an inclination of 25°. The parameters were chosen so that the satellite could be deorbited in a few years. For the orbital inclination angle, 25° was selected as a case where electromagnetic force is highly effective, and 98.4° as a less effective case. At high inclination, very little electromagnetic force is produced as expected, when an electron emitter isn't used. As for the tether width, 150 mm and 25 mm were selected to compare the case where the influence of atmospheric drag is large and the case where it is relatively small. As shown in Figs. 9–12, the effect of atmospheric drag clearly becomes more remarkable as the orbital altitude decreases. Although atmospheric drag is dominant at an orbital inclination angle of 98.4°, the ratio of electromagnetic force increases dramatically at an orbital inclination angle of 25°. According to the above results and previous research [20], the EDT system is less effective in a high inclination orbit where the direction of the geomagnetic field is not suitable. Hereafter, we will introduce an example where an orbital inclination angle is 98.4° as a less effective for an EDT application.
3. Evaluation of the effects on EDT performance by various parameters The effects of orbital conditions and various tether parameters on EDT system performance were investigated by numerical analysis using a simulation code developed by JAXA [18]. To consider tether flexibility, the tether is modeled as a lumped mass by dividing it into a series of point masses, each connected to its neighbors by a spring and a viscous damper (Fig. 6). The equation of motion for each point mass is formulated in a coordinate system wherein the origin is the system's center of mass in LEO (Fig. 7). Orbital perturbations caused by air drag and geopotential are considered using Gauss's vibrational equations of motion. The Naval Research Laboratory's Mass Spectrometer and Incoherent Scatter radar Extended model (NRLMSISE-00) is used for the atmosphere model, the International Reference Ionosphere (IRI) 2016 model is used for the plasma model, and International Geomagnetic Reference Field (IGRF-12) (10 × 10) is used for Earth's geopotential field. F10.7 solar flux and Ap magnetic indices as of 1999 are used, as one example. For electron collection by the bare tether, the two-dimensional orbital motion limit (OML) equation described below was used [19].
enp 2H dJ = dx
2e (x ) me
3.2. Effects on maximum emission current of electron emitter The effects on the maximum emission current of the electron emitter were evaluated. Fig. 13 shows the electromagnetic force and atmospheric drag for each maximum emission current of the electron emitter. The tether width is 25 mm and the tether length is 900 m. It shows the case at an orbital altitude of 800 km and an orbital inclination angle of 98.4°. This result suggests that, with an electron emitter, an EDT system can significantly enhance electromagnetic force.
(1)
In Eq. (1), J (x ), (x ), e, np, H , and me denote the current, potential, elementary electron charge, plasma density, tether width, and electron mass, respectively. The potential of tether is determined by iterating calculations to balance between collected and emitted electrons. The atmospheric drag of the tether depends on the projected area viewed from the orbital velocity direction. Because the attitude around the axis of the tape type tether cannot be controlled, the average projected area is used. As the thickness of the tape type tether is negligible with respect to the width, the average projected area is 2/ of the tether width. In the simulation, the tether is assumed to be aluminum with a thickness of 6 μm, and attached to one side of the ODEX ribbon (Fig. 8). The
Fig. 4. Schematic drawing of the carbon-nanotube-based FEC. 3
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3.3. Effects of tether width and length The effects of tether length and width on performance were evaluated when an emitter is not used. Fig. 14 shows the electromagnetic force and atmospheric drag for each tether width and length. It shows the case at an orbital altitude of 800 km and at an orbital inclination angle of 98.4° without an electron emitter. As shown in Fig. 14, the electromagnetic force is almost proportional to the tether width and proportional to the square of the tether length. And atmospheric drag is also proportional to the projected area. The effect of the tether length is thus considered to be greater than the effect of the tether width. Consequently, the larger the potential difference due to induced electromotive force with a long tether length, and the larger the amount of flowing current. 3.4. Effect of tether conductivity The effect of tether conductivity on electromagnetic force was evaluated. By evaluating the effect of tether conductivity, it is possible to estimate the required thickness of the tether's conductive layer. Fig. 15 shows the electromagnetic force for each tether conductivity. Here, cases of using and not using an electron emitter are assumed. An electron emitter is used with a maximum emission current of 40 mA. It shows the case at an orbital altitude of 800 km and at an orbital inclination angle of 98.4°. The tether width is 25 mm and the tether length is 900 m. As shown in Fig. 15, the electromagnetic force is saturated at a certain conductivity or more. This is because flowing current does not increase even when conductivity is good due to other conditions such as the potential difference. When an electron emitter is used, the saturation point moves to the low resistance side.
Fig. 5. Example of the laboratory model FEC.
3.5. Summary of simulation results and image of deorbit Figs. 16 and 17 show images of the changes in the orbital altitude of the satellite under several conditions relative to the effects of tether length and width, respectively with and without an electron emitter. The initial orbital altitude is 800 km and the orbit inclination angle is 98.4°. As shown in Figs. 16 and 17, it is more effective to extend the tether length than to extend the tether width, and performance can be further improved by using an electron emitter. Even if an orbital inclination angle is 98.4° as a less effective to an EDT application, the 25year rule can be sufficiently satisfied without electron emitter. Fig. 18 shows an image of the change in orbital altitude of the satellite under several inclination angles. As shown in Fig. 18, the EDT system is clearly under a challenging condition at an orbit inclination angle of 98.4°. Therefore, a somewhat elongated tether can be expected to achieve higher performance. However, an elongated tether tends to require a larger storage volume, and a balance between storage volume and performance is important. By making such evaluations, the specifications of the tether and electron emitter can be determined for such EDT requirements as orbital conditions, weight, and deorbit time.
Fig. 6. Lumped mass model.
4. Trial manufacture of tethers and various tests Several EDT prototypes were made to measure the actual tether parameters. The EDT must have specific strength, conductivity, survivability, flexibility, non-stickiness in space, and meet other requirements. Fig. 19 shows several prototype EDTs. The aluminum-deposited polyester film and aluminum-deposited polyimide film are very lightweight, but have poor conductivity due to the deposition of aluminum. Therefore, the thickness of the vapor deposition layer must be increased more than that in general vapor deposition. Given the aluminum surface of the tether, sticking in space may occur, and the surface will become hot due to the poor thermooptical properties of aluminum. Consequently, surface treatment is required.
Fig. 7. The system in LEO.
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Fig. 8. ODEX (silver-plated nylon yarn) ribbon.
Fig. 9. Effects of orbital altitude (tether length: 500 m tether, width: 150 mm, inclination angle: 98.4°).
ig. 10. Effects of orbital altitude (tether length: 900 m, tether width: 25 mm, inclination angle: 98.4°).
Fig. 11. Effects of orbital altitude (tether length: 500 m, tether width: 150 mm, inclination angle: 25°).
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Fig. 12. Effects of orbital altitude (tether length: 900 m, tether width: 25 mm, inclination angle: 25°).
Fig. 13. Effects on maximum emission current of electron emitter.
Fig. 14. Effects of tether width and length.
Fig. 15. Effect of tether conductivity.
The carbon nanotube ribbon is woven in a ribbon shape with carbon nanotube yarn. Carbon nanotube yarn has high specific strength and adequate conductivity. Conductivity can be increased by applying metal plating on the surface of the carbon nanotube yarn. Carbon
nanotube yarn is very flexible compared to carbon fiber. However, carbon nanotube is degraded by atomic oxygen, and thus can be expected to eventually and naturally disappear in space under long-term durations. 6
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Fig. 16. Image of the change in orbital altitude of the satellite (effect of tether length).
Fig. 17. Image of the change in orbital altitude of the satellite (effect of tether width).
Fig. 18. Image of the change in orbital altitude of the satellite (effect of orbit).
Fig. 19. Prototypes of various tethers.
The ODEX ribbon has a good specific strength and is flexible. The ODEX ribbon also has less fuzzing of fibers. But given its low conductivity, measures must be devised to improve conductivity, such as attaching an aluminum foil.
Cracks are considered difficult to develop in the ribbon-like tether, and thus high survivability is expected. The tensile test confirmed how the ribbon-like tether broke. Fig. 20 shows the appearance of the fractured surfaces of the ODEX ribbon and an aluminum-deposited 7
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Fig. 20. Appearance of fractured surfaces.
polyester film subjected to a tensile test for comparison. In the ribbonlike tether, cracks did not spread throughout the tether, and the entire tether did not break. In the future, various tests such as a tensile test and atomic oxygen irradiation test, as well as electrical resistance measurements, will be conducted.
Mitigation Guidelines, IADC Action Item number 22.4 (2007). [5] T. Watanabe, T. Sato, H. Sakamoto, S. Kawamoto, Y. Ohkawa, K. Kamachi, L. Okajima, A study on post-mission disposal device for microsatellite using electrodynamic tape tether Technology, JAXA Special Publication: Proceedings of the 8th Space Debris Workshop, 2019. [6] H. Sakamoto, K. Arai, K. Umetsu, T. Fujita, T. Sato, T. Watanabe, S. Kawamoto, K. Kamachi, Y. Oikawa, L. Okajima, Experimental study on separation of tape tether deployment module for PMD, JAXA Special Publication: Proceedings of the 8th Space Debris Workshop, 2019. [7] G. Sánchez-Arriaga, X. Chen, Modelling and performance of electrodynamic low work-function tethers with photoemission effects, J. Propul. Power 34 (1) (2018) 213–220. [8] J.K. McTernan, S.G. Bilén, Embodied energy repurposing via energy-harvesting electrodynamic tethers, J. Spacecraft Rockets 54 (4) (2017) 789–795. [9] I.C. Bell III, B.E. Gilchrist, J. McTernan, S.G. Bilén, Investigating miniaturized electrodynamic tethers for picosatellites and femtosatellites, J. Spacecraft Rockets 54 (No. 1) (2017) 55–66. [10] K. Iki, S. Kawamoto, Y. Ohkawa, T. Okumura, K. Kawashima, M. Takai, K. Izawa, K. Matsumoto, S. Suzuki, Y. Katayama, Y. Horikawa, K. Inoue, Expected on-orbit tether deployment dynamics on the KITE mission, Trans. JSASS Aerospace Tech. Japan 14 (2016) ists30 Pr_9-Pr_18. [11] M. Khadija, Results of the STEP-AIRSEDS Electrodynamic Tether Deployment Simulations, AIAA Paper 2002-1119 (2002). [12] A. Francesconi, C. Giacomuzzo, F. Branz, E.C. Lorenzini, Survavibility to hypervelocity impacts of electrodynamic tape tethers for deorbiting spacecraft in LEO, 6th European Conference on Space Debris, European Space Operations Centre, Darmstadt, Germany, April 2013, pp. 22–25. [13] K. Makihara, S. Kondo, Structural evaluation for electrodynamic tape tethers against hypervelocity space debris impacts, J. Spacecraft Rockets 55 (4) (January 2018) 1–11. [14] A.H. Fujii, T. Watanabe, H. Sahara, H. Kojima, S. Takehara, Y. Yamagiwa, et al., Space Demonstration of Bare Electrodynamics Tape-Tether Technology on the Sounding Rocket S520-25, AIAA Paper 2011-06503 (2011). [15] Y. Ohkawa, A. Izawa, Y. Yamagiwa, S. Kawamoto, S. Nishida, S. Kitamura, Research and development of carbon nanotube cathodes for electric propulsion, Trans. JSASS Aerospace Tech. Japan 8 (2010) Pb_27–Pb_32 ists27. [16] Y. Ohkawa, K. Iki, T. Okumura, S. Kawamoto, Y. Horikawa, K. Inoue, Y. Kobayashi, T. Uchiyama, T. Kasai, A quick review of an electrodynamic tether experiment on the H-II transfer vehicle, 35th Int'l Electric Propulsion Conf., Atlanta, USA, IEPC2017-359, 2017. [17] Y. Ohkawa, T. Okumura, K. Iki, H. Okamoto, S. Kawamoto, Operation of a carbon nanotube field-emission cathode in low Earth orbit, J. Vac. Sci. Technol. B 37 (2019). [18] S. Kawamoto, Y. Ohkawa, T. Okumura, K. Iki, H. Okamoto, Performance of electrodynamic tether system for debris deorbiting: Re-evaluation based on the results of KITE experiments, IAC-18- A6.6.5, 69th International Astronautical Congress (IAC), Bremen, Germany, October 2018, pp. 1–5. [19] J.R. Sanmartin, M. Martinez-Sanchez, E. Ahedo, Bare wire anodes for electrodynamic tethers, J. Propul. Power 9 (1993) Number 3. [20] T. Sato, S. Kawamoto, T. Watanabe, K. Kamachi, Y. Ohkawa, Y. Oikawa, L. Okajima, H. Okubo, A study on PMD device for microsatellites using electrodynamic tether, 2019-r-29, Proceedings of the 32nd International Symposium on Space Technology and Science, Fukui, Japan, 2019, pp. 15–21 (June).
4. Conclusion We focused on an EDT for microsatellite PMD devices, and considered the material and configuration of the tether. It was assumed that the tape type tether is installed in a folded state and deployed on orbit. By using the folded tape type tether, easy deployment could be expected by ejecting an end mass attached at the end of the tether. And since the device can be miniaturized, it is suitable for microsatellites. The effects of tether parameters were evaluated by numerical simulations. The specifications of the tether and electron emitter could be determined for such EDT requirements as orbital conditions, weight, and deorbit time. We also manufactured several tethers made of new materials based on the simulation results, and tensile tests confirmed how the tethers broke. As a result, the manner of breaking was confirmed as being different among those tethers. In the future, additional ground experiments and numerical analysis will be conducted for a more detailed study. Acknowledgements Part of this research was conducted based on the JAXA Space Innovation through Partnership and Co-creation (J-SPARC) agreement between JAXA and ALE, and the joint research agreement between the Kanagawa Institute of Technology and ALE. References [1] J.-C. Liou, The near-Earth orbital debris problem and the challenges for environment remediation, The 3rd International Space World Conference, 2012. [2] C. Colombo, A. Rossi, F.D. Vedova, A. Francesconi, C. Bombardelli, M. Trisolini, J.L. Gonzalo, P.D. Lizia, C. Giacomuzzo, S.B. Khan, R. G-Pelayo, V. Braun, B.B. Virgili, H. Krag, Effects of passive de-orbiting through drag and solar sails and electrodynamic tethers on the space debris environment, IAC-18-A6.2.8, 69th International Astronautical Congress (IAC), Bremen, Germany, 1-5 October, 2018. [3] R.P. Hoyt, I.M. Barnes, N.R. Voronka, J.T. Slostad, The terminator Tape™: a costeffective de-orbit module for end-of-life disposal of LEO satellites, AIAA SPACE 2009 Conference & Exposition, Pasadena, California, 14-17 September, 2009. [4] Inter-Agency Space Debris Coordination Committee (IADC), IADC Space Debris
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