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Electric power module for a bare electrodynamic tether
Acta Astronautica xxx (xxxx) xxx–xxx
Contents lists available at ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Research paper
Electric power module for a bare electrodynamic tether José A. Carrascoa,b,∗, Francisco García de Quirósa,b, Higinio Alavésa,b, Moisés Navalóna a b
Embedded Instruments and Systems S.L., Avda. de La Universidad Sn, 03202, Elche, Spain Universidad Miguel Hernández, Avda. de La Universidad Sn, 03202, Elche, Spain
ARTICLE INFO
ABSTRACT
Keywords: Electrodynamic tether Space power generation Energy harvesting Spacecraft propulsion de-orbiting
Electrodynamic tethers are proposed as propulsion and energy harvesters for space probes orbiting planets with a magnetic field and ionosphere, however there are no descriptions in the technical literature of the design of an electrical power system for such an application at subsystem and circuit detail. This paper presents a proposal for such a power system that extracts energy from the kinetic energy of a spacecraft using a bare electrodynamic tether, i.e. an unsheathed conductive wire or band, in low-Earth orbit. The application of the system is the powering of the spacecraft while in its final de-orbiting maneuvers at end-of-life with no reliance on the main spacecraft bus.
1. Introduction
architectures with flight heritage in European-based space platforms.
An electrodynamic tether is typically an unsheathed conductive wire, several hundred of meters long, unreeled from a spacecraft orbiting a planet that holds a magnetic field and an ionosphere. The interaction of the free electrons in the wire with the magnetic field follows the basic equation of the Lorenz force [1] and provides either propulsion for orbiting or sling-shot maneuvers [2] or even electrical power to the spacecraft [3]. With growing concerns about the increase of space junk and the incoming regulations to move Earth-orbiting spacecrafts out of their nominal orbits, and even de-orbit them, electrodynamic tethers have been proposed as thruster mechanisms for endof-life orbital operations [4–7], and related projects have been funded by organizations such as the European Commission [8], ESA, NASA and JAXA among others. To make the de-orbiting system completely autonomous from the spacecraft main power bus, energy harvesters have been studied to provide electrical power for mission-related tasks, while the tether is in operation [9]. However, these electrical power systems could be used as main power systems in space missions to planets with very strong magnetic fields and ionospheres such as Earth or Jupiter [10]. As no detailed proposal for such a power system has been found in the literature, this paper makes such a proposal based on electrical
2. Power system proposal
∗
Only a (roughly) detailed power system proposal has been found in the literature [11] and is complemented by the proposal realized herewith. The power system is built under the assumption that the power is delivered by a 20 km conductive tether in a 400 km low-Earth orbit in an average solar activity day with a characteristic I-V curve as shown in Fig. 1. Although this I-V curve is highly dependent on the particular tether implementation and features, such as material, length and geometry, and the orbital height and inclination of the supporting spacecraft, some representative I-V points are represented in Fig. 1 as extracted from Ref. [6]. The variations in intensity and voltage of this curve are due to environmental variations of the spacecraft through its orbit. The power system has three main objectives: 1) provide power to the hollow cathode [12] system, 2) harvest or dissipate (dump) the electrical energy injected by the electrodynamic tether, and 3) use some of it for powering purposes such as supplying the onboard computer that controls the de-orbiting, charge the batteries and balance the spacecraft charge. The block diagram of the complete power supply is presented in
Corresponding author. Embedded Instruments and Systems S.L., Avda. de La Universidad Sn, 03202, Elche, Spain. E-mail address: [email protected] (J.A. Carrasco).
https://doi.org/10.1016/j.actaastro.2019.12.001 Received 28 July 2019; Received in revised form 4 November 2019; Accepted 1 December 2019 0094-5765/ © 2019 IAA. Published by Elsevier Ltd. All rights reserved.
Please cite this article as: José A. Carrasco, et al., Acta Astronautica, https://doi.org/10.1016/j.actaastro.2019.12.001
Acta Astronautica xxx (xxxx) xxx–xxx
J.A. Carrasco, et al. 9
tether and control of the de-orbiting in its early stages. Fig. 3 shows a more detailed block diagram of the proposed electrical power system. The power system in Fig. 3 handles the tether injected energy and powers the hollow cathode during a de-orbiting maneuver. It has to be designed to allow autonomous operation of the complete tether–cathode system once de-orbiting is started. The possibilities for the configuration of such an autonomous system are endless, however, some time ago it became clear that decentralized power systems result in safer and more stable space operation [13,14]. Fig. 3 follows the proposal in Fig. 2 by elaborating a completely decentralized batteryregulated topology with all the different subsystems attached to a common DC battery bus. With such a system it is possible to control the impedances connected to the bus [15] and readily respond to the expected fast transients of the electrodynamic tether. The main components of the power system are:
Current
8
Current (A)
7
6
5
4
3 600
700
800
900
1000
1100
1200
1300
• A charged battery to achieve autonomous operation of the system.
1400
Potential (V)
Fig. 1. The characteristic curve of a 20 km space tether in a 400 km height lowEarth Orbit.
•
Fig. 2. The tether current is directed to a hollow cathode, whose main role is to establish a contact with the surrounding plasma to balance the spacecraft platform charge by releasing electrons to free space by using a jet of ionized argon as described in Ref. [12]. A dissipative load and a main DC/DC power supply are used to condition some of the incoming power to the telemetry and telecommand units of the system and the hollow cathode support subsystems. The system in Fig. 2 is connected to the spacecraft onboard computer, and its power bus, although it would be possible to design the unit to operate autonomously. Nevertheless, even for autonomous operation, the de-orbiting system would need of a secondary battery (i.e. rechargeable) that has to be maintained in good health up to the end of life of the spacecraft to operate the unreel of the
• • •
When the system is in standby, the battery is maintained in trickle charge by using a battery charge regulator (BCR) that extracts power from the spacecraft distribution bus [15]. A BCR DC/DC converter that powers the DC bus while the system is not in de-orbiting operation (standby) to keep the battery in good health by charging it, and to provide telemetry and telecommand (TM/TC) to the de-orbiting system to start its operations at end-oflife. A DC/DC resonant converter [16] that powers the hollow cathode with different voltages from the battery bus of the system. A main DC/DC converter that gets the energy from the electrodynamic tether and injects it into the battery bus to be used for battery charging and powering the hollow cathode. A dump resistor, Rdump , that dissipates any excess energy introduced in the spacecraft by the tether while in operation.
Fig. 2. Block diagram of the power system.
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Fig. 3. Detailed subsystems diagram of the power system.
Fig. 4. The three-domain error amplifier and the definition of the domains of operation.
• A battery error amplifier, BEA, that produces an output proportional
There are four working modes in the system: one occurs when the spacecraft is in operation, and thus the tether system is in standby, and the other three when the spacecraft is being decommissioned by using the electrodynamic tether. The decision of being in standby or de-orbiting is handled by the TM/TC in coordination with the spacecraft onboard computer (OBC). While in standby, the main error amplifier
to the state of charge of the battery, and the main error amplifier that controls the operation of the BEA, which together with a power control sets the voltage at the battery bus and modulates the effect of the Rdump . The power main error amplifier implemenys a technique known as three-domain control as presented in Ref. [17].
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Fig. 5. The isolation barriers and grounding of the power system.
operates the DC/DC BCR for maintaining the battery fully charged. While de-orbiting, the main error amplifier voltage controls in which of the three operational domains is the power system working in and what subsystem is handling the flowing power, Fig. 4 [17]:
operation. While in orbit the power system ensures that the battery remains at maximum load (to be able to respond upon an emergency event) and charges it, when needed, by using a DC/DC converter that acts as a battery charge regulator. While in de-orbiting operation the system control makes the complete system work in one of the following states:
• Battery discharge domain: the tether is not yielding enough energy • •
to power the hollow cathode and charge the battery, which provides such supply from its energy reservoir. Battery charge domain: the battery is charging from the energy provided by the tether, which supplies power to the hollow cathode as well. Dump domain: the tether provides more energy than the needed to power the hollow cathode and charge the battery, which may even become fully charged.
• Battery powers the hollow cathode and control electronics. • Tether powers the hollow cathode and control electronics. • Tether charges the battery in addition to powering the system. • Tether energy exceeds de-orbiting system demand, and excess energy dissipates in the dump resistor.
The system control contains two main circuits: the error amplifier that maintains the operation domain of the complete system (effectively maintaining the voltage of the battery bus and the power harvesting from the tether) as explained in Ref. [19], and the digital circuitry that maintains (and controls) the TM/TC and communicates (if necessary) with the spacecraft. The system control also handles the following functionalities:
The main error amplifier in Fig. 4 implements a three-domain controller [17] that makes use of a proportional-integral-derivative (PID) amplifier to define, by comparing the battery bus voltage and a reference voltage, an error that is minimized and used to govern the power flow within the bus. These schemes are widely used in space systems and have been proposed in other applications [18]. Due to the flight heritage that may be found in the referred literature, usual techniques for reliability and single-point failure-free (SPFF) design may be followed to produce a spacecraft flight hardware using the described topology. The control block is in charge of the proper operation of the complete system during the functioning in orbit and during the de-orbiting
• Communications with spacecraft OBC through RS422, or equivalent, interface. • Analog to digital conversion of telemetry parameters. • Generation of references to power supplies for the subsystem. • Storage of spacecraft parameters is the units is expected to survive de-orbiting.
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supplies are galvanically isolated from the spacecraft main power bus and the battery bus. 3. The input DC/DC converter power stage design This converter has the structure of a standard battery regulator. The requirement for a it is to convert the DC bus voltage to a lower voltage (i.e. the battery bus voltage) while controlling the current injected from the bus. A Buck type regulator has been selected, due to simplicity of design, over the low-ripple Buck, which has been mostly used in the past [14,20]. Further, the classical Buck is preferred since the input attenuation of the input filter is far greater than that of the low-ripple Buck. The main characteristics of this subsystem are:
• Hot redundancy: two Buck converters working in parallel. • Single-point failure-free power supplies within each converter. • Proven MOSFET drivers [21]. 4. Battery dimensioning The selection of a +42 V Li–Ion battery is a reasonable state of the art technology for space batteries. This provides up to 1 h of autonomous operation for the de-orbit system, with no external power input, which seems a rather safe figure considering a delivered power of around 250 Wh. The battery may be implemented by two strings in parallel of ten battery cells of type MP176065, manufactured by SAFT, in series, which is technically known as a 10s2p configuration. The weight of the complete package is less than 2 kg. Fig. 6 shows a diagram of the battery. 5. Tether power conditioning The tether power conditioning needs a converter that transforms a voltage from up to 1500 V to the battery bus voltage. For protection purposes, i.e. proper grounding definition, we will need a converter with galvanic separation between the tether and the power distribution bus. The power conditioner has two parts: a step-down regulator that keeps the voltage regulated at approximately 50 V across it and a one switch zero voltage zero current (ZVZC) [16] power converter in series with it, see Fig. 7. The design of the step-down regulator is straightforward as it is a classical converter. The important consideration in this design is the possibility of having the entire voltage produced by the tether across its main switch, which would require the used of high voltage switches that do not exist for space applications. To withstand this voltage the solution proposed in Ref. [22] has been followed with the MOSFET transistor IRFP23N50 (that has the space-qualified equivalent JAXA R 25K4188). The control of this stage is realized with the PWM controller UC1823. Fig. 6. Schematic view of the battery of the system and protections.
6. Dumping resistor power conditioner The dumping resistors have been designed in order to dissipate power from the tether and control the current through it, which in turn affects the drag force. In order to have a versatile control of such power a parallel connection of 10 switchable resistors has been planned. Further, each of these resistors is treated as a separate item that may be configured later in response to specific mission requirements. Such an arrangement is desirable because it provides system redundancy as well. The switching of the resistors is governed following the logic of an
A fundamental aspect of the whole system is the connection of the spacecraft chassis to the electrical reference and its relation to the plasma. Fig. 5 shows the location of the tether reference. Every block has implemented a galvanic isolation (i.e. electrical isolation by using appropriate transformers). At every moment, the tether is isolated from the spacecraft and shares reference with the hollow cathode, while the system provides a path for current flow; besides, the hollow cathode
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Fig. 7. The tether conditioning DC/DC main power supply.
array of sequential switching shunt regulators (S3R) conditioning a solar array [19]. The block diagram of five dump resistors is presented in Fig. 8. Each resistor string is switched on and off by the main error amplifier output voltage as presented in Fig. 4. As for the design we have manufactured the resistor as an integrated track on a standard Eurocard printed circuit board, to minimize volume, with the following electrical requirements:
• Power: 1 kW. • Resistance: 1 k . • PCB Dimension: 100 × 160 mm • Copper thickness: 100 μm. • Clearance: 100 μm.
2
resistor is implemented as 663 segments of 150 mm length over the two sides of the PCB. The maximum dissipation density is 5.9 W per mm3. An aluminum panel located at the bottom of each PCB is used to transfer heat toward the spacecraft chassis. 7. Hollow cathode power conditioner Space heritage and authors’ experience for implementing a multiple output power supply for conditioning a hollow cathode with the specifications in Ref. [23] suggests the use of a zero voltage zero current switching (ZVZCS) unregulated push-pull transformer [16] with a Buck (i.e. step down) pre-regulator. Being the ZVZCS converter not regulated at its output, the use of a Buck pre-regulator requires that it handles the entire power, which may decrease the efficiency of the complete power converter by 5%. However, the very high efficiency of the ZVZCS for the power levels considered (around 98%), results in a combined regulated power supply with a final efficiency better than 93%. Fig. 10 shows the a simplified schematics of the ZVZCS converter, implemented with only one switch (i.e. half ZVZCS converter) and the outputs that supply the hollow cathode subsystems.
Eurocard standard size.
The track length for the desired resistance is obtained as follows:
R=
l , a
(1)
where ρ is the copper resistivity, l the length of the conductor and a its cross section area. By using this equation we may implement a 100 m resistor as a copper track 100 μm in height and 17 μm in width with a 100 μm clearance between them in the PCB as represented in Fig. 9. The
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Fig. 8. Tether conditioning with five similar strings of dump resistors. (FPD stands for FET/MOSFET pulse driver.)
redundancies that will require duplications for the subsystems that incorporate heavy components like inductors, transformers, or power capacitors. Only the dumping resistors, that are implemented with a very simple circuit, are implemented in cold redundancy, which will help to handle and distribute the heat generated by them. The whole system operation and coordination is controlled by the main error amplifier as described in Refs. [24,25]. In addition to single-point failure-free circuits and redundancies, we also consider over-current and over-voltage protections. 9. Conclusion The specific electronics at system, subsystem, and circuit level for conditioning a tether at low-Earth orbit for the application of de-orbiting has not been described in the technical literature thus far. In this paper, a complete power system for such an application is proposed at system level and its implementation at subsystem and circuit level is described. The final application of such a system is the autonomous assistance in the de-orbiting of spacecraft at end-of-life, an optional requirement for spacecraft in low-Earth Orbits at the writing of this paper that will likely be mandatory in the new future.
Fig. 9. Dump resistor PCB dimensioning.
8. System protections In order not to increment the size and the mass of the complete power system, it is reasonable to think of the implementation of singlepoint failure-free circuits and not to implement hot or cold
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Fig. 10. Hollow cathode DC/DC power supply.
Acknowledgments
[12] J.D. Williams, P.J. Wilbur, Ground-based tests of hollow cathode plasma contactors, 3rd International Conference on Tethers in Space - toward Flight, 1989, pp. 77–87. [13] A. Capel, D. O'Sullivan, High power conditioning for space applications, ESA SPS 91 - Power from Space, 1991, pp. 437–446. [14] A. Capel, D. O'Sullivan, J.C. Marpinard, High-power conditioning for space applications, Proc. IEEE 76 (4) (1988), https://doi.org/10.1109/5.4425 391–0408. [15] ECSS-E-ST-20C Space Engineering: Electrical and Electronic, European Space Agency, 31 July 2008. [16] A.H. Weinberg, L. Ghislanzoni, A new zero voltage and zero current powerswitching technique, IEEE Trans. Power Electron. 7 (4) (1992) 655–665, https:// doi.org/10.1109/63.163645. [17] D. O'Sullivan, Space power electronics-design drivers, ESA J. 18 (1994) 1–23. [18] J. Carrasco, E. Dede, J. Benavent, F. Bordry, A. Dupaquier, A. Ferreres, A three domain controller for a high frequency high power four quadrant power converter for superconducting magnets, PESC Record - IEEE Annual Power Electronics Specialists Conference, vol. 1, 1997, pp. 356–362, , https://doi.org/10.1109/PESC. 1997.616749. [19] D. O'Sullivan, A. Weinberg, The sequential switching shunt regulator (S3R), Proceedings of the Third ESTEC Spacecraft Power Conditioning Seminar, ESA SP126, 1977. [20] K. van Dijk, J. Klaassens, H. Spruijt, D. O'Sullivan, Battery charger design for the columbus MTFF power system, IEEE Trans. Aerosp. Electron. Syst. 33 (1997) 29–37, https://doi.org/10.1109/7.570705. [21] A. Crausaz, E. Gasquet, Power mosfet: Esa driving license, ESA SP-369, Fourth European Space Power Conference, vol. 1, 1995, pp. 227–233. [22] H.L. Hess, R.J. Baker, Transformerless capacitive coupling of gate signals for series operation of power mos devices, IEEE Trans. Power Electron. 15 (5) (2000) 923–930, https://doi.org/10.1109/63.867682. [23] K. Xie, R.A. Martinez, J.D. Williams, Current-voltage characteristics of a cathodic plasma contactor with discharge chamber for application in electrodynamic tether propulsion, J. Phys. D Appl. Phys. 47 (15) (Apr. 2014), https://doi.org/10.1088/ 0022-3727/47/15/155501. [24] W. Knorr, Power system of meteosat second generation, ESA SP-416, Proceedings of the Fifth European Space Power Conference, 1998, pp. 11–16. [25] J.E. Haines, D. Levins, A. Robben, A. Sepers, The meteosat second generation (msg) power system, IECEC 97 Proceedings of the Thirty Second Intersociety Energy Conversion Engineering Conference (Cat. No. 97CH6203), vol. 1, 1997, pp. 538–543, , https://doi.org/10.1109/IECEC.1997.659247.
This work has been funded by the European Commission under the FP7/Space Project 262972. The authors thank the company Embedded Instruments and Systems S.L. for its support on the realization of this project. References [1] R. P. Feynman, The Feynman Lectures on Physics: Millennium Ed., vol. 2, Basic Books. [2] D. L. Gallagher, L. Johnson, J. Moore, F. Bagenal, Electrodynamic Tether Propulsion and Power Generation at Jupiter, NASA/TP-1998- 208475 (01 1998). [3] S. Bilén, J. Mcternan, B. Gilchrist, I. Bell, N. Voronka, R. P Hoyt, Electrodynamic tethers for energy harvesting and propulsion on space platforms, AIAA SPACE Conference and Exposition, 2010, https://doi.org/10.2514/6.2010-8844. [4] E. Ahedo, J. Sanmartin, Analysis of electrodynamic tethers as deorbiting systems, 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2000, https://doi.org/10.2514/6.2000-3763. [5] C. Bombardelli, J. Herrera Montojo, J. Pelaez, A. Iturri-Torrea, Space debris removal with bare electrodynamic tethers, AAS/AIAA Space Flight Mechanics Meeting, vol. 136, 2010, pp. 2523–2534. [6] C. Bombardelli, J.M. Nuñez-Ayllón, J. Peláez Álvarez, Performance analysis of bare electrodynamic tethers as microsats deorbiting systems, AAS/AIAA Space Flight Mechanics Meeting, vol. 136, 2010, pp. 2515–2522. [7] M. Sanjurjo Rivo, J. Pelaez, Energy analysis of bare electrodynamic tethers, J. Propuls. Power 27 (1) (2011) 246–256, https://doi.org/10.2514/1.48168. [8] Thin as a Rail, Strong as a rock, Research EU Results Magazine No. 32, (2014). [9] C. Bombardelli, Power density of a bare electrodynamic tether generator, American Institute of Aeronautics and Astronautics (AIAA), vol. 28, 2012, https://doi.org/10. 2514/1.b34189. [10] C. Bombardelli, E. Lorenzini, J.R. Sanmartin, Jupiter power generation with electrodynamic tethers at constant orbital energy, J. Propuls. Power 25 (2009) 415–423, https://doi.org/10.2514/1.38764. [11] Ball Aerospace Systems Division, Electrodynamic Tether System Study : Final Report, National Aeronautics and Space Administration, 1987.
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Contents lists available at ScienceDirect
Acta Astronautica journal homepage: www.elsevier.com/locate/actaastro
Research paper
Electric power module for a bare electrodynamic tether José A. Carrascoa,b,∗, Francisco García de Quirósa,b, Higinio Alavésa,b, Moisés Navalóna a b
Embedded Instruments and Systems S.L., Avda. de La Universidad Sn, 03202, Elche, Spain Universidad Miguel Hernández, Avda. de La Universidad Sn, 03202, Elche, Spain
ARTICLE INFO
ABSTRACT
Keywords: Electrodynamic tether Space power generation Energy harvesting Spacecraft propulsion de-orbiting
Electrodynamic tethers are proposed as propulsion and energy harvesters for space probes orbiting planets with a magnetic field and ionosphere, however there are no descriptions in the technical literature of the design of an electrical power system for such an application at subsystem and circuit detail. This paper presents a proposal for such a power system that extracts energy from the kinetic energy of a spacecraft using a bare electrodynamic tether, i.e. an unsheathed conductive wire or band, in low-Earth orbit. The application of the system is the powering of the spacecraft while in its final de-orbiting maneuvers at end-of-life with no reliance on the main spacecraft bus.
1. Introduction
architectures with flight heritage in European-based space platforms.
An electrodynamic tether is typically an unsheathed conductive wire, several hundred of meters long, unreeled from a spacecraft orbiting a planet that holds a magnetic field and an ionosphere. The interaction of the free electrons in the wire with the magnetic field follows the basic equation of the Lorenz force [1] and provides either propulsion for orbiting or sling-shot maneuvers [2] or even electrical power to the spacecraft [3]. With growing concerns about the increase of space junk and the incoming regulations to move Earth-orbiting spacecrafts out of their nominal orbits, and even de-orbit them, electrodynamic tethers have been proposed as thruster mechanisms for endof-life orbital operations [4–7], and related projects have been funded by organizations such as the European Commission [8], ESA, NASA and JAXA among others. To make the de-orbiting system completely autonomous from the spacecraft main power bus, energy harvesters have been studied to provide electrical power for mission-related tasks, while the tether is in operation [9]. However, these electrical power systems could be used as main power systems in space missions to planets with very strong magnetic fields and ionospheres such as Earth or Jupiter [10]. As no detailed proposal for such a power system has been found in the literature, this paper makes such a proposal based on electrical
2. Power system proposal
∗
Only a (roughly) detailed power system proposal has been found in the literature [11] and is complemented by the proposal realized herewith. The power system is built under the assumption that the power is delivered by a 20 km conductive tether in a 400 km low-Earth orbit in an average solar activity day with a characteristic I-V curve as shown in Fig. 1. Although this I-V curve is highly dependent on the particular tether implementation and features, such as material, length and geometry, and the orbital height and inclination of the supporting spacecraft, some representative I-V points are represented in Fig. 1 as extracted from Ref. [6]. The variations in intensity and voltage of this curve are due to environmental variations of the spacecraft through its orbit. The power system has three main objectives: 1) provide power to the hollow cathode [12] system, 2) harvest or dissipate (dump) the electrical energy injected by the electrodynamic tether, and 3) use some of it for powering purposes such as supplying the onboard computer that controls the de-orbiting, charge the batteries and balance the spacecraft charge. The block diagram of the complete power supply is presented in
Corresponding author. Embedded Instruments and Systems S.L., Avda. de La Universidad Sn, 03202, Elche, Spain. E-mail address: [email protected] (J.A. Carrasco).
https://doi.org/10.1016/j.actaastro.2019.12.001 Received 28 July 2019; Received in revised form 4 November 2019; Accepted 1 December 2019 0094-5765/ © 2019 IAA. Published by Elsevier Ltd. All rights reserved.
Please cite this article as: José A. Carrasco, et al., Acta Astronautica, https://doi.org/10.1016/j.actaastro.2019.12.001
Acta Astronautica xxx (xxxx) xxx–xxx
J.A. Carrasco, et al. 9
tether and control of the de-orbiting in its early stages. Fig. 3 shows a more detailed block diagram of the proposed electrical power system. The power system in Fig. 3 handles the tether injected energy and powers the hollow cathode during a de-orbiting maneuver. It has to be designed to allow autonomous operation of the complete tether–cathode system once de-orbiting is started. The possibilities for the configuration of such an autonomous system are endless, however, some time ago it became clear that decentralized power systems result in safer and more stable space operation [13,14]. Fig. 3 follows the proposal in Fig. 2 by elaborating a completely decentralized batteryregulated topology with all the different subsystems attached to a common DC battery bus. With such a system it is possible to control the impedances connected to the bus [15] and readily respond to the expected fast transients of the electrodynamic tether. The main components of the power system are:
Current
8
Current (A)
7
6
5
4
3 600
700
800
900
1000
1100
1200
1300
• A charged battery to achieve autonomous operation of the system.
1400
Potential (V)
Fig. 1. The characteristic curve of a 20 km space tether in a 400 km height lowEarth Orbit.
•
Fig. 2. The tether current is directed to a hollow cathode, whose main role is to establish a contact with the surrounding plasma to balance the spacecraft platform charge by releasing electrons to free space by using a jet of ionized argon as described in Ref. [12]. A dissipative load and a main DC/DC power supply are used to condition some of the incoming power to the telemetry and telecommand units of the system and the hollow cathode support subsystems. The system in Fig. 2 is connected to the spacecraft onboard computer, and its power bus, although it would be possible to design the unit to operate autonomously. Nevertheless, even for autonomous operation, the de-orbiting system would need of a secondary battery (i.e. rechargeable) that has to be maintained in good health up to the end of life of the spacecraft to operate the unreel of the
• • •
When the system is in standby, the battery is maintained in trickle charge by using a battery charge regulator (BCR) that extracts power from the spacecraft distribution bus [15]. A BCR DC/DC converter that powers the DC bus while the system is not in de-orbiting operation (standby) to keep the battery in good health by charging it, and to provide telemetry and telecommand (TM/TC) to the de-orbiting system to start its operations at end-oflife. A DC/DC resonant converter [16] that powers the hollow cathode with different voltages from the battery bus of the system. A main DC/DC converter that gets the energy from the electrodynamic tether and injects it into the battery bus to be used for battery charging and powering the hollow cathode. A dump resistor, Rdump , that dissipates any excess energy introduced in the spacecraft by the tether while in operation.
Fig. 2. Block diagram of the power system.
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Fig. 3. Detailed subsystems diagram of the power system.
Fig. 4. The three-domain error amplifier and the definition of the domains of operation.
• A battery error amplifier, BEA, that produces an output proportional
There are four working modes in the system: one occurs when the spacecraft is in operation, and thus the tether system is in standby, and the other three when the spacecraft is being decommissioned by using the electrodynamic tether. The decision of being in standby or de-orbiting is handled by the TM/TC in coordination with the spacecraft onboard computer (OBC). While in standby, the main error amplifier
to the state of charge of the battery, and the main error amplifier that controls the operation of the BEA, which together with a power control sets the voltage at the battery bus and modulates the effect of the Rdump . The power main error amplifier implemenys a technique known as three-domain control as presented in Ref. [17].
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Fig. 5. The isolation barriers and grounding of the power system.
operates the DC/DC BCR for maintaining the battery fully charged. While de-orbiting, the main error amplifier voltage controls in which of the three operational domains is the power system working in and what subsystem is handling the flowing power, Fig. 4 [17]:
operation. While in orbit the power system ensures that the battery remains at maximum load (to be able to respond upon an emergency event) and charges it, when needed, by using a DC/DC converter that acts as a battery charge regulator. While in de-orbiting operation the system control makes the complete system work in one of the following states:
• Battery discharge domain: the tether is not yielding enough energy • •
to power the hollow cathode and charge the battery, which provides such supply from its energy reservoir. Battery charge domain: the battery is charging from the energy provided by the tether, which supplies power to the hollow cathode as well. Dump domain: the tether provides more energy than the needed to power the hollow cathode and charge the battery, which may even become fully charged.
• Battery powers the hollow cathode and control electronics. • Tether powers the hollow cathode and control electronics. • Tether charges the battery in addition to powering the system. • Tether energy exceeds de-orbiting system demand, and excess energy dissipates in the dump resistor.
The system control contains two main circuits: the error amplifier that maintains the operation domain of the complete system (effectively maintaining the voltage of the battery bus and the power harvesting from the tether) as explained in Ref. [19], and the digital circuitry that maintains (and controls) the TM/TC and communicates (if necessary) with the spacecraft. The system control also handles the following functionalities:
The main error amplifier in Fig. 4 implements a three-domain controller [17] that makes use of a proportional-integral-derivative (PID) amplifier to define, by comparing the battery bus voltage and a reference voltage, an error that is minimized and used to govern the power flow within the bus. These schemes are widely used in space systems and have been proposed in other applications [18]. Due to the flight heritage that may be found in the referred literature, usual techniques for reliability and single-point failure-free (SPFF) design may be followed to produce a spacecraft flight hardware using the described topology. The control block is in charge of the proper operation of the complete system during the functioning in orbit and during the de-orbiting
• Communications with spacecraft OBC through RS422, or equivalent, interface. • Analog to digital conversion of telemetry parameters. • Generation of references to power supplies for the subsystem. • Storage of spacecraft parameters is the units is expected to survive de-orbiting.
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supplies are galvanically isolated from the spacecraft main power bus and the battery bus. 3. The input DC/DC converter power stage design This converter has the structure of a standard battery regulator. The requirement for a it is to convert the DC bus voltage to a lower voltage (i.e. the battery bus voltage) while controlling the current injected from the bus. A Buck type regulator has been selected, due to simplicity of design, over the low-ripple Buck, which has been mostly used in the past [14,20]. Further, the classical Buck is preferred since the input attenuation of the input filter is far greater than that of the low-ripple Buck. The main characteristics of this subsystem are:
• Hot redundancy: two Buck converters working in parallel. • Single-point failure-free power supplies within each converter. • Proven MOSFET drivers [21]. 4. Battery dimensioning The selection of a +42 V Li–Ion battery is a reasonable state of the art technology for space batteries. This provides up to 1 h of autonomous operation for the de-orbit system, with no external power input, which seems a rather safe figure considering a delivered power of around 250 Wh. The battery may be implemented by two strings in parallel of ten battery cells of type MP176065, manufactured by SAFT, in series, which is technically known as a 10s2p configuration. The weight of the complete package is less than 2 kg. Fig. 6 shows a diagram of the battery. 5. Tether power conditioning The tether power conditioning needs a converter that transforms a voltage from up to 1500 V to the battery bus voltage. For protection purposes, i.e. proper grounding definition, we will need a converter with galvanic separation between the tether and the power distribution bus. The power conditioner has two parts: a step-down regulator that keeps the voltage regulated at approximately 50 V across it and a one switch zero voltage zero current (ZVZC) [16] power converter in series with it, see Fig. 7. The design of the step-down regulator is straightforward as it is a classical converter. The important consideration in this design is the possibility of having the entire voltage produced by the tether across its main switch, which would require the used of high voltage switches that do not exist for space applications. To withstand this voltage the solution proposed in Ref. [22] has been followed with the MOSFET transistor IRFP23N50 (that has the space-qualified equivalent JAXA R 25K4188). The control of this stage is realized with the PWM controller UC1823. Fig. 6. Schematic view of the battery of the system and protections.
6. Dumping resistor power conditioner The dumping resistors have been designed in order to dissipate power from the tether and control the current through it, which in turn affects the drag force. In order to have a versatile control of such power a parallel connection of 10 switchable resistors has been planned. Further, each of these resistors is treated as a separate item that may be configured later in response to specific mission requirements. Such an arrangement is desirable because it provides system redundancy as well. The switching of the resistors is governed following the logic of an
A fundamental aspect of the whole system is the connection of the spacecraft chassis to the electrical reference and its relation to the plasma. Fig. 5 shows the location of the tether reference. Every block has implemented a galvanic isolation (i.e. electrical isolation by using appropriate transformers). At every moment, the tether is isolated from the spacecraft and shares reference with the hollow cathode, while the system provides a path for current flow; besides, the hollow cathode
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Fig. 7. The tether conditioning DC/DC main power supply.
array of sequential switching shunt regulators (S3R) conditioning a solar array [19]. The block diagram of five dump resistors is presented in Fig. 8. Each resistor string is switched on and off by the main error amplifier output voltage as presented in Fig. 4. As for the design we have manufactured the resistor as an integrated track on a standard Eurocard printed circuit board, to minimize volume, with the following electrical requirements:
• Power: 1 kW. • Resistance: 1 k . • PCB Dimension: 100 × 160 mm • Copper thickness: 100 μm. • Clearance: 100 μm.
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resistor is implemented as 663 segments of 150 mm length over the two sides of the PCB. The maximum dissipation density is 5.9 W per mm3. An aluminum panel located at the bottom of each PCB is used to transfer heat toward the spacecraft chassis. 7. Hollow cathode power conditioner Space heritage and authors’ experience for implementing a multiple output power supply for conditioning a hollow cathode with the specifications in Ref. [23] suggests the use of a zero voltage zero current switching (ZVZCS) unregulated push-pull transformer [16] with a Buck (i.e. step down) pre-regulator. Being the ZVZCS converter not regulated at its output, the use of a Buck pre-regulator requires that it handles the entire power, which may decrease the efficiency of the complete power converter by 5%. However, the very high efficiency of the ZVZCS for the power levels considered (around 98%), results in a combined regulated power supply with a final efficiency better than 93%. Fig. 10 shows the a simplified schematics of the ZVZCS converter, implemented with only one switch (i.e. half ZVZCS converter) and the outputs that supply the hollow cathode subsystems.
Eurocard standard size.
The track length for the desired resistance is obtained as follows:
R=
l , a
(1)
where ρ is the copper resistivity, l the length of the conductor and a its cross section area. By using this equation we may implement a 100 m resistor as a copper track 100 μm in height and 17 μm in width with a 100 μm clearance between them in the PCB as represented in Fig. 9. The
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Fig. 8. Tether conditioning with five similar strings of dump resistors. (FPD stands for FET/MOSFET pulse driver.)
redundancies that will require duplications for the subsystems that incorporate heavy components like inductors, transformers, or power capacitors. Only the dumping resistors, that are implemented with a very simple circuit, are implemented in cold redundancy, which will help to handle and distribute the heat generated by them. The whole system operation and coordination is controlled by the main error amplifier as described in Refs. [24,25]. In addition to single-point failure-free circuits and redundancies, we also consider over-current and over-voltage protections. 9. Conclusion The specific electronics at system, subsystem, and circuit level for conditioning a tether at low-Earth orbit for the application of de-orbiting has not been described in the technical literature thus far. In this paper, a complete power system for such an application is proposed at system level and its implementation at subsystem and circuit level is described. The final application of such a system is the autonomous assistance in the de-orbiting of spacecraft at end-of-life, an optional requirement for spacecraft in low-Earth Orbits at the writing of this paper that will likely be mandatory in the new future.
Fig. 9. Dump resistor PCB dimensioning.
8. System protections In order not to increment the size and the mass of the complete power system, it is reasonable to think of the implementation of singlepoint failure-free circuits and not to implement hot or cold
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Fig. 10. Hollow cathode DC/DC power supply.
Acknowledgments
[12] J.D. Williams, P.J. Wilbur, Ground-based tests of hollow cathode plasma contactors, 3rd International Conference on Tethers in Space - toward Flight, 1989, pp. 77–87. [13] A. Capel, D. O'Sullivan, High power conditioning for space applications, ESA SPS 91 - Power from Space, 1991, pp. 437–446. [14] A. Capel, D. O'Sullivan, J.C. Marpinard, High-power conditioning for space applications, Proc. IEEE 76 (4) (1988), https://doi.org/10.1109/5.4425 391–0408. [15] ECSS-E-ST-20C Space Engineering: Electrical and Electronic, European Space Agency, 31 July 2008. [16] A.H. Weinberg, L. Ghislanzoni, A new zero voltage and zero current powerswitching technique, IEEE Trans. Power Electron. 7 (4) (1992) 655–665, https:// doi.org/10.1109/63.163645. [17] D. O'Sullivan, Space power electronics-design drivers, ESA J. 18 (1994) 1–23. [18] J. Carrasco, E. Dede, J. Benavent, F. Bordry, A. Dupaquier, A. Ferreres, A three domain controller for a high frequency high power four quadrant power converter for superconducting magnets, PESC Record - IEEE Annual Power Electronics Specialists Conference, vol. 1, 1997, pp. 356–362, , https://doi.org/10.1109/PESC. 1997.616749. [19] D. O'Sullivan, A. Weinberg, The sequential switching shunt regulator (S3R), Proceedings of the Third ESTEC Spacecraft Power Conditioning Seminar, ESA SP126, 1977. [20] K. van Dijk, J. Klaassens, H. Spruijt, D. O'Sullivan, Battery charger design for the columbus MTFF power system, IEEE Trans. Aerosp. Electron. Syst. 33 (1997) 29–37, https://doi.org/10.1109/7.570705. [21] A. Crausaz, E. Gasquet, Power mosfet: Esa driving license, ESA SP-369, Fourth European Space Power Conference, vol. 1, 1995, pp. 227–233. [22] H.L. Hess, R.J. Baker, Transformerless capacitive coupling of gate signals for series operation of power mos devices, IEEE Trans. Power Electron. 15 (5) (2000) 923–930, https://doi.org/10.1109/63.867682. [23] K. Xie, R.A. Martinez, J.D. Williams, Current-voltage characteristics of a cathodic plasma contactor with discharge chamber for application in electrodynamic tether propulsion, J. Phys. D Appl. Phys. 47 (15) (Apr. 2014), https://doi.org/10.1088/ 0022-3727/47/15/155501. [24] W. Knorr, Power system of meteosat second generation, ESA SP-416, Proceedings of the Fifth European Space Power Conference, 1998, pp. 11–16. [25] J.E. Haines, D. Levins, A. Robben, A. Sepers, The meteosat second generation (msg) power system, IECEC 97 Proceedings of the Thirty Second Intersociety Energy Conversion Engineering Conference (Cat. No. 97CH6203), vol. 1, 1997, pp. 538–543, , https://doi.org/10.1109/IECEC.1997.659247.
This work has been funded by the European Commission under the FP7/Space Project 262972. The authors thank the company Embedded Instruments and Systems S.L. for its support on the realization of this project. References [1] R. P. Feynman, The Feynman Lectures on Physics: Millennium Ed., vol. 2, Basic Books. [2] D. L. Gallagher, L. Johnson, J. Moore, F. Bagenal, Electrodynamic Tether Propulsion and Power Generation at Jupiter, NASA/TP-1998- 208475 (01 1998). [3] S. Bilén, J. Mcternan, B. Gilchrist, I. Bell, N. Voronka, R. P Hoyt, Electrodynamic tethers for energy harvesting and propulsion on space platforms, AIAA SPACE Conference and Exposition, 2010, https://doi.org/10.2514/6.2010-8844. [4] E. Ahedo, J. Sanmartin, Analysis of electrodynamic tethers as deorbiting systems, 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, 2000, https://doi.org/10.2514/6.2000-3763. [5] C. Bombardelli, J. Herrera Montojo, J. Pelaez, A. Iturri-Torrea, Space debris removal with bare electrodynamic tethers, AAS/AIAA Space Flight Mechanics Meeting, vol. 136, 2010, pp. 2523–2534. [6] C. Bombardelli, J.M. Nuñez-Ayllón, J. Peláez Álvarez, Performance analysis of bare electrodynamic tethers as microsats deorbiting systems, AAS/AIAA Space Flight Mechanics Meeting, vol. 136, 2010, pp. 2515–2522. [7] M. Sanjurjo Rivo, J. Pelaez, Energy analysis of bare electrodynamic tethers, J. Propuls. Power 27 (1) (2011) 246–256, https://doi.org/10.2514/1.48168. [8] Thin as a Rail, Strong as a rock, Research EU Results Magazine No. 32, (2014). [9] C. Bombardelli, Power density of a bare electrodynamic tether generator, American Institute of Aeronautics and Astronautics (AIAA), vol. 28, 2012, https://doi.org/10. 2514/1.b34189. [10] C. Bombardelli, E. Lorenzini, J.R. Sanmartin, Jupiter power generation with electrodynamic tethers at constant orbital energy, J. Propuls. Power 25 (2009) 415–423, https://doi.org/10.2514/1.38764. [11] Ball Aerospace Systems Division, Electrodynamic Tether System Study : Final Report, National Aeronautics and Space Administration, 1987.
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