Survey
* Your assessment is very important for improving the workof artificial intelligence, which forms the content of this project
* Your assessment is very important for improving the workof artificial intelligence, which forms the content of this project
De-orbiting Spacecraft with Electrodynamic Tether Devices Carmen Pardini IADC AI 19.1 on “Benefits and Risks of Using Tethers in Space” Space Flight Dynamics Laboratory ISTI/CNR, Via G. Moruzzi 1 56124 Pisa, ITALY 21st IADC Meeting, 10-13 March 2003, Bangalore, INDIA Introduction Electrodynamic tether drag can provide a cost-effective method for autonomously de-orbiting low earth orbit (LEO) spacecraft to mitigate the growth of orbital debris De-orbiting devices based on the use of conducting tethers have been recently proposed as innovative solutions to remove satellites and upper stages from low earth orbit once they have completed their missions Studies of such devices are currently being planned, or are in the early development phase in the US and Europe A flight experiment to validate the performance of the bare electrodynamic tether in space and demonstrate its capability to produce thrust is scheduled by NASA at the end of March 2003 Electrodynamic Drag Concept The concept of the electrodynamic tethers for de-orbiting applications is based on the exploitation of the Lorentz’s force due to the interaction between the electric current flowing in the conductive tether and the geomagnetic field The decelerating Lorentz force (electrodynamic drag) depends in a complex way on the design parameters of the system, the orbit and the characteristics of the local ionosphere Fdrag I (l ) dl B L where dl is the differential element of tether length, B the local magnetic field and I(l) the current flowing in the wire The mechanical power dissipated by the drag force can be expressed as P Fdrag v0 I (l ) ( v0 B ) dl L where v0 is the velocity vector of the system. Because the drag and the power are proportional to the product L I, it is better to generate high currents and minimize the tether length. Shorter tethers translate into lighter systems as well as a decrease in the probability of collisions with other systems orbiting at LEO heights De-orbiting Time [1] The time t needed to de-orbit a spacecraft between two given heights (corresponding to orbital radii a1 and a2 with a1< a2 ) is given by (Vannaroni et al.) a2 G M e mt da 2 2 a Fdrag v0 a1 t where G is the gravitational constant (G = 6.673 x 10-20 km3 kg-1 s-2), Me is the mass of the earth (Me = 5.973 x 1024 kg), mt is the spacecraft mass including the tether system and the far-end-mass, v0 is the orbital velocity and a is the orbital radius. Thus the de-orbiting time can be calculated once the drag force has been computed as a function of the orbit altitude The decay rate is accelerated in correspondence of the lower heights for the larger currents in the tether due to the higher density of the ionospheric plasma. De-orbiting Time [2] The maximum efficiency is obtained for equatorial orbits, due to a combination of larger induced voltages and ionospheric densities For nearly polar orbits the tether’s interaction with the geomagnetic field is much lower and de-orbiting times with low-mass electrodynamic tethers are rather high Another important parameter to consider is the tether length, its value determines the induced voltage and therefore the current. Short tethers imply significantly longer decay times, due to the combining effect of smaller currents and induced voltages However, although the performances of long tethers are attractive, the price to pay in terms of mass, risk of arching and debris impact could be too high for reliable operations Electrodynamic Space Tethers Proposed to De-orbit Spacecraft The Electrodynamic De-Orbiting And Re-entry Device (EDOARD) is jointly developed in Italy by Alenia Spazio and the University of Rome “La Sapienza” in view of potential commercial exploitation. EDOARD is designed to de-orbit satellites and upper stages in the 600 to 4000 kg mass range In between 600 and 2000 km of orbital altitude and up to an orbital inclination of 65° Tether Unlimited Inc. (USA) is developing a lightweight, reliable space tether system called the Terminator TetherTM (TT) to remove defunct satellites from low earth orbit. It represents a commercialized version of the ProSEDS experiment and is currently built to provide de-orbit capability for a 2000-3000 kg LEO spacecraft EDOARD The EDOARD system is intended to provide the carrier spacecraft with an electrodynamic device capable to de-orbit it within a few months The EDOARD mass is envisaged to be less than 30-35 kg (between 1% and 5% of the carrier vehicle mass at launch) The EDOARD system is based on a short conductive tether (4-5 km long), mechanisms capable of autonomously deploying the tether from the carrier vehicle after a ground command, electrical/mechanical subsystems assuring adequate current collection and emission and the associate control electronics The electron-collecting tether end is equipped with a large inflatable passive electron collector (up to 10 m diameter), which increases the efficiency of the system while reducing the tether length The tether system stabilization control will be actively performed by the EDOARD electrical subsystem during the orbital decay phase EDOARD Schematic Configuration L. Iess et al., Acta Astronautica Vol. 50, No. 7, pp. 407–416, 2002 Terminator TetherTM The Terminator TetherTM (TT) is a small, lightweight, low-cost device that will be attached to satellites and upper stages before launch. It will be composed of a conducting, survivable tether, a tether deployment system, a device for emitting electron current, and an electronic control system (TCU) Tether: to electrically insulate the host spacecraft from the tether, a short section of the tether near the spacecraft will be constructed of high-strength, nonconducting yarns. The rest of the tether will be a survivable HoytetherTM structure of thin aluminum or copper wires. The tether design will vary upon the mass and orbit of the host spacecraft. For a typical LEO satellite massing 1500 kg, the tether will be 5 km long and mass ~ 15 kg. The HoytetherTM design will enable the tether to provide a very high probability of surviving the orbital debris environment for the period of several weeks or months required to de-orbit a spacecraft Ground control signals can be sent to the TCU to perform avoidance maneuvers The Terminator Tetherª. Copyright © 1999 by Tethers Unlimited, Inc. Published by the American Institute of Aeronautics and Astronautics with permission. Flight Demonstration of Electrodynamic Drag ProSEDS The idea of using electrodynamic drag to remove spacecraft from orbit was first discussed by Joseph P. Loftus of NASA/JSC in June 1996. The Loftus electrodynamic drag de-orbit concept will be demonstrated with the flight of the Propulsive Small Expendable Deployer System (ProSEDS) schedule at the end of March 2003. ProSEDS will be the first mission to produce electrodynamic thrust, use a bare wire tether, and recharge batteries using tethergenerated power ProSEDS will be carried into a ~ 360-km circular orbit as a secondary payload on an Air Force Delta II rocket, its main cargo being a Global Positioning System satellite The experiment will attempt to de-orbit the second stage of Delta (mass ~ 900 kg) with a 5 km bare aluminium tether connected with a 10 km nonconductive tether. The overall mass of the tether system is 100 kg inclusive of tether, deployer, batteries, plasma contactor and scientific instrumentation to measure the system performances The decay rate estimated with a numerical code developed at the Harvard-Smithsonian Center for Astrophysics is about 20 km/day and the re-entry time is ~ 15 days. How ProSEDS Works http://astp.msfc.nasa.gov/proseds/images/tech_whole_large.jpg De-orbit Times Using Prototypes of the EDOARD Tethered System [1] Altitude Interval [km] De-orbiting Time [days] 1500 – 1400 37.99 1400 – 1300 40.91 1300 – 1200 46.02 1200 – 1100 40.18 1100 – 1000 35.06 1000 – 900 39.08 900 – 800 33.97 800 – 700 32.87 700 – 600 25.93 600 – 500 20.09 500 – 400 16.07 400 – 300 13.15 300 – 200 10.96 1500 – 200 392.28 Initial Altitude Orbital Inclination 1500 km 55 Impedance 280 Payload Mass 2000 kg Tether Length 5 km Electron Collector Diameter 10 m Year of Mission 2003 De-orbit Times Using Prototypes of the EDOARD Tethered System [2] for a typical satellite of 500 kg mass L. Iess et al., Acta Astronautica Vol. 50, No. 7, pp. 407–416, 2002 De-orbit Times Using a Model of the Terminator TetherTM [1] The TetherSim numerical model was used to determine the rate at which a Terminator TetherTM system can de-orbit a spacecraft A Terminator Tether massing ~ 2% of the mass of the host spacecraft could de-orbit an upper stage from a 400 km, 50° orbit within 2 weeks a communication satellite from a 850 km, 50º orbit within ~ 3 months a SkyBridge satellite from a 1475 km, 55° orbit in ~ 1.2 years A lightweight TT system could effectively de-orbit spacecraft with inclinations up to 75° In the following figures, the tether system modeled consisted of a 15 kg aluminum tether, with a 15 kg endmass. The tether was chosen to be 7.5 km long. The host spacecraft mass was 1500 kg, so the TT system massed 2% of the host satellite mass The results revealed that the rate of descent varies approximately as the cosine of the orbital inclination De-orbit Times Using a Model of the Terminator TetherTM [2] http://www.tethers.com/papers/TTReno00.pdf De-orbit Times Using a Model of the Terminator TetherTM [3] http://www.tethers.com/papers/TTReno00.pdf Electrodynamic Tether System (EDTS) by Delta-Utec De-orbit times have been computed with ETBsim The descent rate of a 700 kg satellite, de-orbited with an EDTS, as a function of altitude was computed for an equatorial orbit. The tether length was assumed to contain an additional 3 km mechanical tether (1 kg), the EDTS sub-satellite was 15 kg and the EDT diameter was 0.8 mm 60 7.5 km elec tether 12.5 km elec tether Descent rate [km/day] 50 40 30 20 10 0 300 500 700 900 1100 1300 1500 Altitude [km] E. Van der Heide & M. Kruijff, Acta Astronautica Vol. 48, No. 5-12, pp. 503-516, 2001 Conclusions For typical electrodynamic tether lengths of 5-10 km, the Lorentz force reduces the mean altitude of the tethered system orbit at rates from 2 to 50 km/day, decreasing with increasing debris mass, inclination and altitude At altitudes > 2000 km the plasma density and magnetic field strength are insufficient and electrodynamic tethers are inefficient to de-orbit spacecraft The performances of the previous ED tethers have been investigated for nearly circular orbits. In fact, for highly eccentric orbits, such as GTO, a tether system is no longer stable or librating, but will start to rotate at pass of perigee. A full damping of the rotation will require a major design challenge