Transcript De-orbiting Time - PUMA

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