Tethers and debris mitigation

~

Acta Astronautica Vol. 48, No. 5-12, pp. 503-516, 2001 © 2001 Deha-Utec SRC Pubhshed by Elsevier Science Ltd Printed in Great Britain

Pergamon

www.elsevier.com/locate/actaastro

PII: S0094-5765(01 )00074- I

0094-5765/01 $- see front matter

Tethers and debris mitigation Erik Jan van der Heide & Michiel Kruijff DeltaJJtec SRC, Leiden, The Netherlands,www.de~ta-utec.com

Abstract

Although tethers may have large exposure in terms of areatime product, they deliver a quick cleaning service that may be appreciated by the future users of space. © 2001 Delta-Utec SRC. Published by Elsevier Science Ltd.

in recent years, the use of tethers has been proposed for reduction of space debris either through momentum transfer or use of electrodynamic effects. Tethers have been shown to at least theoretically allow for quick, elegant and cost-effective deorbit of defunct satellites or spent stages. On the other hand, the large risk that tethers themselves may pose to other satellites in orbit has been recognized as well, The large collision area of tethers, combined with operational hazards and meteoroid risk may result in a large orbital exposure. For example, in 1997, the ESA/Dutch 35-kin tether deployment of YES from TEAMSAT was inhibited after an analysis of the collision risk for the case the tether operation would fail. The question rises how these two points of view compare to eachother. This paper intends to highlight a representative selection of the proposed tether applications while taking into account the added risks caused by the tethers themselves. Typical applications fi'om recent literature will be briefly described, such as an Ariane 502 spent stage re-entry fi'om GTO and the concept of deboost of defunct satellites by interaction of a conductive tether with the Earth magnetic field. Mass savings of the tethered systems versus conventional equivalents will be evaluated. Based on a crude risk analysis, involving elements such as mission complexity, dynamic stability, meteoroid risk and orbital life time, a general outline of limiting factors can be given for the various applications. Special attention is reserved for implementation of mechanisms that help reduce this tether risk, such as the DUtether (Tether Degradable by Lrltraviolet), utilization of airdrag and solar pressure, the effect of residual current in bare tethers, tether retrieval etc. It is proposed how a net tether-induced mitigation can be compared to that of conventional alternatives, i.e. deboost by rocket engine or a completely passive approach. This comparison is put in the perspective of an everincreasing occupation of the space environment. It is concluded that tethers can in fact help mitigate the debris risk and that for each application a useful niche can be defined. It is argued that eliminating pollution directly after use of the precious resource of space is not only good custom, but also an important way to make the risk of debris controllable and independent of future trends.

1.

Introduction

Space debris is becoming an issue, a well-known fact. Not so well understood is the interaction between space tethers and the debris environment. Some people discuss the potential risks that tethers may pose, while others stress the usefulness of tethers for debris mitigation. Because of this ambiguous information it is hard to judge the balance between the two ext,emes. In fact, there are many interactions and factors that do not allow one to have a simple single view. It is attempted to assess the considerations and effects that should be taken into account in discussions about tethers and debris mitigation. We will focus in this paper on possible uses of tethers for debris mitigation and create an unbiased first order measure for their efficiency, safety and effectiveness when compared to conventional alternatives. Target debris ander investigation are: • Spent stages: they are considered to have prime share in catastrophic collisions and the future debris environment ~. • Constellations: because they would occupy a single altitude with large amounts Tether concepts studied are: M o m e n t u m transfer, lightweight mechanical tethers can exchange momentum due to the effect of gravity gradient and/or inducing a swing velocity to the tethered system • Lorentz drag force: a bare tether with cathode at the lower end moving through the magnetic field of the Earth will collect electrons from the plasma and a current is induced. Again due to the Earth's magnetic field, a Lorentz drag force is created. •

General assumptions and constraints for this study, some quantified, some based on the personal point of view of the authors, are discussed in section 2. The next chapter gives an overview of possible methods for tethered debris

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504

51 st L4 F Congress

mitigation. Section 4 describes a method developed to compare the risk of a tether with conventional alternatives. This method is applied in section 5 and section 6 addresses the interpretation of the results.

2.

breakage will avoid a recoil of the endmass and add to the system's overall safety.

2.2 Tether failure and degradable tether development

Considerations for the comparison

I f for some reason the tether deployment or control system

malfunctions, and the tether would become a hazard, or as a last resort to avoid a collision, it is assumed it is possible to disconnect the tether completely on either side by an

We have limited ourselves to the current state of space practice and near term tether applications. in this section, we will further simplify the range of design and contingency possibilities according to state-of-the art technology as well as our current personal view. Analysis is only done in rough order of magnitude.

2.1 Tether design is insensitive and brNk.mffe

autonomous system 2 or telecommand. This w i l l increase

the tether's drag area to mass ratio. The decoupled tether remaining in orbit is eventually re-entered primarily due to the combined mec,hanisan of ak drag, solar pressure I° and ion/electron currents [Discussion Box in Appendix]. For these mechanisms to be sufficiently helpful in degradation of the orbit, limitations to orbit of choice apply [Table I ]. Lunar-solar perturbations are independent of tether drag area and are relatively small disturbances for tethers. The) are periodic and non-dissipative and their main affect is on the perigee altitude of GTO orbits with tens of kms, depending on the constellation of Sun and Moon]L The density increase can have a significant effect on orbital lifetime (factor 2 for 5-10% reduction in altitude).

meteoroid/debris

It is assumed in this paper that tethers that will be flown will not or very rarely be cut by meteoroids or debris. This is to be achieved by faiisafe design (multi-strand tethers such as Carroll Caduceus 2 or Hoytethcr 3) or safe-life design (tapes). The single strand 4 lan x 2 mm TiPS 4 tether is currently in orbit for over 4 years, favoring the analytical and heuristic models tuned with early flight data5"6 that were considered too optimistic by extrapolation of ground based debris-impact tests7.s. Multi-strand tethers may as yet face some disadvantages and unknowns, such as performance under torsional vs. electrostatic forces, reduced break strength over weight ratio 2 and increased deployment friction, but it is assumed that such tethers become commonplace for any multi-week mission and will be developed further into perfection. Break strength should be scaled to possible tension jerks (during deployment or operation) and/or centripetal force 9. For LEO, a reasonable safety factor with respect to maximum deplo)Tnent tension is 4. On top of this, a safety factor on tether strength is required with respect to the

A possibility to extend the use of mechanical tethers to virtually any orbit, is the DUTether, or Tether Degradable by Ultraviolet 12. Such a tether is currently under development and would degrade into free molecules in a matter of months -compared to a mission time of days-. Full tether failure would then create a reduced risk of order l0 2 in high LEO orbits9 or order 10.3 in GTO, similar to that of the Shuttle's TSS missions 13. However, since we are focussing on current state-of-the-art, for this paper it is assumed that DUTethers cannot yet be used. 2.3 Stability can be ensured Mechanical tethers are rather stable in space and show only minor excitation of out-of plane dynamics, even after long periods of time 4. Due to the inclined magnetic field, conductive tethers endure an out-of-plane force component causing

industrial value. When a Dyneema tether is bent around small radius, its strength may decrease a factor 2-3 2. The safety-factor can easily be chosen such that ifsomehow the tether would get a hard-jam very early in the deployment, the tether will break since there is little stretch capability invoking a high tension peak~. Such planned contingency

Solar pt~mum LEO, i.-45:<800 kin' GTO: pgee < 800 km

I

LEO: <550 krn GTO: pgee <400 km

ConducUve

LEO. <500 km GTO" pgee <400 km

LEO. i--45<700km GTO:pgee< 7001¢n

GTO: pgee <500 km

Free tether acceleration

-0.5 [150 kml -10s[700±150 kin]

~10 ~

LEO, 1<65:<1500km LEO.i<75" <700 loin GTO: pcjee< 600 Im~ -10" [400 km] -10e [1500 kin]

Dependingstrongly on solar c'yde

GTO: 60% reduced launch ~ndow dependingon relative Sun-orbit constellation

GTO passagethrough i o n o s ~ is fast but shod, effect 1 order of magnitudeless

Dependingon relative SunMoon-Orbit constellation

Qenoral (airdrag only)

MQclmnlcal

[,m:!

m

cBrmnt

l.unas-8olar

Combined

GTO: pgee

LE(~ <550 km LEOi--45:<800

<500 km

GTO: pgee <800 km LEO, i<65:<1500 km LEO. i<75 <700 GTO: pgee < 700 Ion

±10~ High pgee GTO launch Umeand date to be selectedfor solar pressure.Aunarsolar effects

Table h Rough upper fimits of application area for typical tethers (2-7 mZ/kg) from contingency case risks (free flying tether lifetime due to several effects) - assuming no DUTetbcrs.

51st L4F Congress

complicated dynamics, including excitation of transversal waves, that are especially difficult to stabilize with aid of simple hardware (tension measurement, rough angular knowledge from accelcrometers or single GPS receiver) t4. The instability increases with inclination and current (max at 300 kin). Because the CtaTent decreases strongly at high inclinations, instability reaches its maximum for medium inclination. Even more challenging effects however are the ionospheric irregularities and the Earth's 436 km magnetic dipole offset. For a typical case (7 Ion I mm bare tether, h=i000 kin, i=50°) ETBSim [section 5. I] shows that after - a day transversal waves are inducing rapidly an instability that does not occur without these effects. The long term stability problem for bare eleetrodynamic tethers has only recently been rec,ognized. Because the problem is very interesting and many academic minds are involved in tether dynamics control, it is expected that a solution will be found shortly, at least to constrain the oscillations below 30 degrees amplitude. It is assumed that this can be done without complicating design. Simple control actuators are (closed loop) current limitation or damping by extra deployment. Use of differential GPS and/or thrusters is considered not competitive. Importantly, current control may have a major impact on the feasible downward rate, especially at low altitudes ~5. For the moment, stability can be increased by using a longer mechanical tether than required just for safe deployment, at the expense of -2 kg tbr 10 km (e.g. ProSEDS) lb. This only has a definite effect if the conductive tether is attached directly to the hea D debrisside rather than to the much lighter balancing mass, as not to increase the moment arm of the Lorentz-force with the mechanical part. Z4 Tether-tether collisions If tethers would be the way to proceed for deorbiting, one could wonder whether they would become so abundant that tether-tether collisions become an issue, clearly an unwelcome scenario. Ifa couple of constellations would be electrodynamically deorbited in overlapping periods of 5 years, about 40 deorbits/year sounds reasonable, equivalent to - 4 tethers in space at the same time. It should be possible to coordinate the deorbits of these to avoid inter-tether collisions.

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and a low mass, low cost and high reliability is required to make them compete with the rocket alternative. Successful tether deployment with very simple lightweight hardware has been elegantly demonstrated in space2°. Very precise deployment with such hardware has been demonstrated recently in ground tests-'L2,~.Deployment of a tether can be done passively (in axial direction) from a spool and is inherently stable, whereas retrieval is much more challenging to control 's and requires a motorized reel, winding positioning hardware and additional state measurement systems. Retrieval of significant tether lengths has not yet been demonstrated. It is therefore assumed that a deployer without retrieval capabilities is used, except for the Travelling Tether concept, for which, because of reusability, the system mass is much less of a problem.

3.1

Electrodynamic Tether Deorbit System A tethered satellite system travelling through the Earth magnetosphere will induce a field, called EMF, due to the Hall-effect or alternatively, induction of the second type. The system's velocity and the magnetic field (strength and direction) result in a voltage drop of the plasma with respect to the tether of some 100 to 200 V per km tether length. This induced field ensures that electrons are collected from the plasma on the bare tether. When no potential difference is applied at the upper satellite to overrule the induced field, a current will flow upwards in the tether, causing a Lorentz drag force [see discussion box]. The electrons emitted at the lower satellite by a plasma contaetor will Iravel down some hundreds of Ion following the magnetic field lines and bounce between ionospheric boundary layers, but never find their way back to the tether. At every reflection some energy is converted and finally they disappear in the background plasma again. The principle has already been demonstrated successfully in the Charge, Oedipus, PMG and TSS experiments. For typical tether lengths of 5-10 Ion and diameter o f - I mm the Lorentz-force is some dN at maximum plasma density (300kin) to mN at high altitudes (1500 km). This force will reduce the semi-major axis (and hence the mean altitude) of the tethered system orbit at rates from 2-50 kin/day, decreasing with increasing debris mass. inclination or altitude. Because of its simplicity and effective current collection, a

3.

Tether concepts for debris mitigation

The basics and features of three possible concepts for debris mitigation are shoal), discussed in this section: • Electrodynamic drag Is'" • Momentum transfer 17'Is • Combination of the above: Travelling tether concept to If tethers are to be used for debris mitigation, for most (early) applications they will become add-ons to satellites

full)' bare aluminum tether 2~'25 is assumed for deorbit purposes. It should be considered that such a tether is rather rigid and

might cause difficulties during initial (low tension) deployment. An initial deployment of a few km polyethylene (Dyneema/Spectra) may serve as the way to increase the gravity gradient to a sufficient level to allow for deployment of the conductive part. A tether needs to have an endmass on either side for safe deplo)raent and stability. Since a competitive deorbit system will be designed for minimum mass, the tether

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50b

deployer, cathode (device for emission of electrons) and electronics itself will serve as endmass. Since the cathode needs to be on the lower side, the tether needs to be deployed downward from the attached piece of debris. The system will thus look like this:

Mechanttal lelher

\ t Bare lethet l ,collects elo:tror, s i

r

(~ Oepto,,el"& ,:alhoOc (emns ¢ k'cffons i

Figure I: Electrodynamie Tether System (EDTS) 3.1 1

EDTS in LEO

EDTS systems have been proposed particularly tbr LEO constellation deorbit, because of the large potential of identical satellites that could be designed to include the EDTS and keep costs log~4. But the s)stem might also be a candidate for spent stages and dedicated satellites. The advantage for spent stages is that the system can be designed for short lifetimes. Also, the required de-orbit fuel mass for spent stages is often considerable [Figure 9]. The EDTS effectiveness however can be easily reasoned to be limited in orbital inclination {low current at high geomagnetic inclination, geomagnetic field inclined b) 11.5°). high altitude (low current and large AV requirement). However, there might also be a limitation at low altitudes. Since the air drag at low altitudes by itself is quite effective in dissipating orbital energy, the remaining gain in total deorbit time that a EDTS will achieve becomes less and less (especially belog 300 km where also the plasma density starts to decrease). Because of the large collision area of the tether, it might e~en be advantageous to disconnect the tether at a certain altitude, leaving its cathode attached [See discussion box]. The ne~ and much lighter tether system will quickly re-enter. The debris might take some tens of years still, but with a very small collision area. How the effectiveness of the EDTS for debris mitigation in LEO compares to alternatives is subject of sections 4.5.6. 3.1.2 EDTS in GTO It is assumed that the EDTS will (initially) only be used in LEO:

At high circular altitudes (>2000 km) the plasma density and magnetic field strength are insufficient. Once could think however of using an electrodynamic tether in GTO to deorbit upper stages2°. For highly eccentric orbits (e>0.4) with low perigee, such as GTO, a tether system is no longer stable or librating, but will start to rotate at pass of perigee ~7, about once per hour. Such rotation will degrade performance, even when a cathode is placed on either side. Full damping of the rotation would require active reeling, a major design challenge. The passage through the dense plasma will be fast but short, only about 20 minutes, whereas the LEO satellite constellation passage lasts about 30 minutes. As the orbital energy that is to be dissipated is O(10) times larger, deorbit time would be large, even when measured by total LEO passage time. Therefore the risk posed by the tether's sweeping volume can be expected to be an order of magnitude larger than for bare tethers connected to LEO debris. As we will see in Section 5, the tethered deorbit ti'om LEO is only improving the sweeping volume within a single order of magnitude at best. So because of both risk and mass considerations an electrodynamic tether in GTO for deorbit will not stand the comparison with conventional rocket deboost.

3.2

Momentum transfer

A totally different concept for dcorbit is that of momentum transfer. This method uses a lightweight Dyneema tether (licensed in the USA as Spectra) to connect and restrain two satellites at different orbital height. The gravi~ gradient over the height difference casu quo the centripetal force will keep the tether tensioned. The satellites are forced to orbit with the orbital angular velocity of the common center of mass. In circular orbiL upon cut of the tether, momentum transfer is effectuated as the upper mass (with a velocity too fast to remain in circular orbit) obtains an increased apogee and the lower mass therefore reduces its perigee. The maximum orbital change that can be achieved expressed in I as distance to the common center of mass is listed in Table 2: Orbit LEO circular LEO circular GTO apogee

Release from tether hanging swinging rotating

APeri~leem, ~ 7I 13 / 40 /

Table 2:Perigee change after tethered momentum transfer

Successful momentum transfer demonstrations to date in LEO have been the SEDS-I. SEDS-2 and ]'SSI-R missions. YES contained a 35 km tether in GTO, but was not deployed because of unfavorable solar pressure. Fhe advantage of deboost by momentum transfer over the EDTS is threefold: • Deplo)ment and libration dynamics are largel.~ in-plane and easy Io t.'onlrol

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51~t L4F Congres~

low LEO debris, if: • such deboost is desired from the launcher side and debris community • the payloads find advantage in the extra orbital height or (perhaps alternati~,ely) mass budget • once the tethcr is used for higher altitudes than recommended in section 2.2 the system is highly reliable and validated in lower orbits

The boost works two ways: as the debris is re-entered, a useful satellite can he raised ]'he tether remains in orbit very shortly (hours-days). if it is designed to sta) attached to the debris. ]he additional collision risk created by the tether can only get significant if the tether itself would end up freely in orbit [see 2.31. The disadvantages are: • No reusability without reeling (whereas Lorentz force can he actively reversed) • Limited AV capability (whereas electrodynmnic tethers can continue to deliver thrust indefinitel20 • Use fbr debris mitigation limited to spent stages (since the) bring the satellite that ~ill be Ix~sted into higher orbit) 'and ISS v,aste. 3 2.1

3.2.2

system I'~. By cutting the rotating system in apogee, the perigee can be brought down into the attnosphere, which is more effective than the eleca-odynamic method [3.1.2] that first circularizes the stage and then spirals it down. Deployment of a tether in GTO has been found to be very straightforward, lengths up to 40 km can be deployed at constant speed of 2 m/s without any form of active braking or control, on the arc towards apogee 2s. 30% Of the current upper stages are a candidate fur this deorbit I. Stage deorbiting is of special interest fur Ariane 5 (13%?), since it has a vet3 high perigee of 620 km (lifetime of hundreds of years) and will be the major contributor to space debris in GTO. Some sample cases are perlbrmed [Table 4], under the same assumptions as in Section 3.2. I. The safety factor on the tether design is 4. Furthermore it should be considered that the normal engine option might require extra batteries/solar cells to make it to apogee (-- 5 hours). This mass is cancelled out in belo,,, cases against the tether electron ics and batteries.

Momentum transfer from cimular orbits

The use of a ~20-40 km swinging tether for ISS waste disposal has been extensively studied and found both

technically and financially feasible, significantly' reducing the propellant required for ISS station-keeping, e.g. 27.21 But removal of spent stages would have the most significant impact on the future debris environment/. Spent stages will often have rockets available (the choice to not re-enter the spent stage is made as result of a trade-off). If re-entry is nevertheless desired, the tether system mass is to be compared with the fuel mass that can be saved. As for rocketed re-entry, an initial attitude is required to facilitate deployment. The margins are large {-20-30 degrees all directions)". lable 3 reports our analysis based on data of spent stagesk Assumed is Is,=300 s, use of a Dyneema swinging tether, de-orbit of the spent stage to 60 km perigee ~, a safety factor on tether strength of 6 [section 2. I], and a tether %stem (TS) mass relation of mrs ,x-mr,~h,,°~, using a SEDS t2,.pe deployer 5 as a starting point. The second iteration is not taken into account (i.e. the fact that mass of the tether and deboost rocket system will decrease the available payload mass and orbit gain). Good candidates are stages that are relatively' light compared to the payload mass, at relatively low altitudes, using relatively short tethers. This is true for 28% of all upper stages. Tethered momentum transfer is thus a good candidate fur reducing Fraction mass

Stage

%

Delta II

Lowest mass rsehna-2 Highest mass ;'end Lowest or1~1 New Bt~ I-t~hest od~t Scaled average

ling]

stqe vn.

1 1 8C 05. 4~

924 8300 1976 4185 11164

1~ 325C 650(; 500~ 3678

Idtitude Tether

[km]

It is tbund that the advantages of a tether are only significant for Ariane 5. As tbr the LEO case, the P/L mass should be as high as possible. Increase of tether length above the minimum required does not lead to significant improvement. Controlled excitation of the rotation (that almost doubles tether pcrtbrmance with respect to the first orbit) can take

3 days or more. It is part of the system trade-off to decide whether 70-q0 kg of extra payload is worth 3 more days of launcher operation plus delay for the payload itself. It can be imagined that the rotation can be purposefull) ApoP4.

lengthIkml Ikm]

1000 850 275 21,251 642

Momentum transfer from GTO

Also spent stages in GTO might be denrbited via a tether

131~ 216 ~.=22 292

1827 2868 340 3853 961

PedP/I.

guelsafo [ko]

[km]

stage 106x 100~, 28( 225~ 8el

pn. 83 638 43 760 113

73 57' 4,E 601 11(

TeIJ~ermass TethersysMm

SF=6)Ikal mssmL[k,tl islhwllml 132 1587 10 2827 87

161 1506 20 2978 111

-1 .-48Z 6~ -1617 11~

Table 3

Stage Anane 5 Single ~,nane 5 Single Anane 5 Double Anane 5 Double Anan'= 4 S,ngle Titan/Centaur HH CZ3-3A

Table 4

'mass Min,mum L Bigger L Min,mum L Bigger L Mimmumr'rL Minimum L Mlrllmum L Mtmmum L

3138 3138

3138 3138 1240 354O 3000 3000

[kg]

Tether mass Tether system )tdvallllge kltitude Tether Apo PIL Ped PIL Fuel safe [kg] [kln] len~lth ~ m ] ['km] [km] susie P/L ;SF:6) [Ikl~ mills e e l [kg] ~ l . , r llm] 68OO 62C 20 35869 87£ 62 3 60 21 28 94 5000 62 3 144 123 146 60 62C .5( 35879 12~ 1900 62C 37 35886 154E 62 3 58 37 48 72, 1900 62 3 78 68 84 57 82c 16= 2O0( 207 35995 ZcJ~ 65 8 1 2 1~ 45OO 207 7 35996 32~ 18 6 22 2 4 37 2200 207 .¢ 35¢Jg6 407 158 17 2 4 2~. 1000 207 1E 36004 648 15 8 16 4 6 2E

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enhanced or dampened with electrodynamic tethers to improve momentum transfer, but a few Newton thrust during up to 10% ofthe orbit can only deliver several m/s (or I-2 km tether), and this at the cost of a more complex system.

3.3

Travelling tether concept

A third option is the travelling tether that uses both electrodynamic drag and momentum transfer. It is a system that would travel through space using electrodynamic drag to initialize deorbit of one piece of debris. It would finally realease it into the atmosphere and use the momentum transfer and electrodynamic propulsion to maneuver itself close to the next piece of debris. And so on. It is not a major problem if such a system is very heavy because of reusability. Differential GPS and a reeling system would be required, as well as a long wide tether to limit deorbit

rather complex, it will not likely be a near term application and theretbre falls somewhat out of the scope of this paper. Nevertheless, a first assessment from the debris mitigation's point of view can be made. In the TERESA concept ~ it is claimed that -100 objects could be removed in a timespan of 7 years using a 93 km tether. So deorbit times per satellite are comparable to the EDTS, but the tether is 10 times longer. If the tether would be continuously exposed during this period, the sweep volume would be l0 times as large. It should therefore be attempted to significantly limit the tether exposure, leaving a major challenge to quickly adjust inclination 3°.

4.

A risk comparison method

How can the debris mitigation performance of alternative systems be compared.'? An tmbiased comparison of tethered deorbit scenarios with alternatives is important, difficult,

RISK CALCULA TJONPARAMETERS

Note

Measurl

LEGEND

max I

avoid pet" 3 or t0 /days # avoidances

collision

Ilotlv e xvoidant:4 IWotlme by iorllele¢ current

risk as 2.2. but

r~k ml. s m a |

li

Option P Process (multiplloatlve) - - -IP Possible smn|rlos (additive) Calculation input Secondal~/ order

Table 5" Method schematic. L = length of tether, S -- averaged width of safety, area around an operational satellite. The values are first order guesses. See main text for explanation

51st IAF Congrers system mass, reliability and cost, the actual risk of collision imposed by the solution of choice is of course the major parameter that needs to be evaluated. This can be expressed with the help of a scaled sweeping volume, rather than lifetime or a tether-area-lifetime product 3~ as will be explained in this section. In order to get to comparable risk esthnates, the sweeping volume is to be multiplied by the debris density for the appropriate debris size range. A method has been developed that breaks up the risk into 3 components, that can then be individually compared:

1. 2. 3.

Expected number of tether maneuvers to avoid an operational satellite Expected number of avoidance maneuvers required by operational satellites Risk of break-up collision

ad I.If the tether system is designed to avoid satellites, it should be operationall) viable (tether length and rather large alert boxes around satellites yields large interception area). One could think of avoiding debris/unguided S/C as well. when operationally viable. Unless avoidance turns out to be a trivial matter the collision risk is seen as the preferred measure as it involves a much smaller interception area. ad 2.1ftbe tether system descent rate is uncontrolled it would be at the least annoying, and perhaps expensive, when man)' operational satellites need to avoid the tether system. On the other hand, the operators of satellites should show some understanding lbr the good cause of debris reduction. It is thus fair to compare ~ith BAU the estimate of the total number of required avoidance maneuvers during the total lit~ime. ad3.A break-up collision, no matter what the secondary consequences may be. is assumed to be catastrophic b) itself. No further risk assessment is performed on the parts after break-up. This category also includes the operational satellites that are not able to acti~el) a~oid debris. The method is schematicall.,, presented in Figure 2 and explained in the following sections. Business as usual (no active de-orbit, BAU) and de-orbit with autonomous rocket boost are compared to EDTS deorbiting in order to create a measure for the mitigation effectiveness of the tether deboost. 1"he sweep area is the averaged projected area around the debris (÷ tether) under investigation, through which operational satellites or other debris should not pass. This area moves at orbital speed during the degradation time. The product of the averaged values is the sweeping volume, from which the risk can be determined:

Risk = f pobjeas V~er = f Pobj~.=s, 4 ~

vo,bit~¢d~radatio.

The factor [ is used to take into account the averaged relative speed effect with the debris and the projected future increase of the debris ensironment. Since the latter

509

effect is assumed of predominant influence on the relative results for the alternatives, we d u b f t h e Future Factor. Future trends for operational satellites should be taken into account to the max. Much work has been performed on predicting the futm'e space use='H. Expectations are an almost linear increase with about factor 3 uncertainty. If we are to take preventive measures now, as not to ruin the space environment in the far future, we should do this also for the case that future use of space will be intensive. Therefore we should akeady now prepare for a possible strong increase in use of operational satellites. Considering the number of spent stages and debris, a more moderate increase can be considered than the worst case exponential scenario. On one side BAU is a hypothetical case where we don't seem to care about keeping space clean. On the other hand future users can be expected to take more and more responsibility to carry de-orbit systems or launch stages into reduced-lifetime orbits. In this paper, also for simplicity, a linear Increase of both debris and operational satellites with a factor 3 in 100 years is assumed. Propagation of risks over degradation time is limited to 150 years, as we find it not reasonable to propagate for hundreds of years. If anything can be expected about such. time periods, it will be that we have a completely different vies of or approach to space and getting into space. At some point we will have such simple access to space and/or such perfect automated collision avoidance mechanisms that today's debris will be another problem solved. Or perhaps we will not go into LEO anymore at all. The choice of 150 years is rather arbitrary but serves as an example to create awareness of the lack of meaning of indefinite risk propagation.

4.1

Autonomous rocket deboost

The risk for a conventional, but autonomous rocket deboost is orders of magnitude smaller than for the BAU case and governed solely by its reliability. Operations are minimal and deboost is fast and relatively accurate. If the boost fails, the system would end up in the BAU configuration. Since the latter possibility is similarly present for the EDTS option both options cancel eachother out in the comparison.

4.2

Business as usual (BALI)

4.2.1 Avoidance by operational SIC Sweeping area is determined by the safety envelope around such spacecraft (lO0s of meters to several kilometerslW). Such envelopes are typically several times longer (velocity, vector direction) than wide. In the future the safety envelops can be expected to go down, because of technical advances and practical/cost considerations. An average of 100 m is assumed.

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51st IAF Congrcss

4.2.2

Break-up collisions

The chance of break-ups should be considered for debris and unguided operational spacecraft. A break-up or catastrophic event is simplified to be a collision with debris energy > 40 kJ/kg, tbr a typical case this is > 10 cm I ~. The collision area (sweeping area) is defined by the square of the dead satellite/spent stage's typical dimension (2.57 m satellite, 3.1 m spent stage1).

4.3

De-orbit by EDTS

A deorbit by EDTS has several failure scenarios. A failure before deplo)ment: does not affeet the comparison between the several scenarios [see 4. I 1. Considering avoidance of operational S/C. two choices can be made: controlled or uncontrolled de-orbit. The risk of break-up collision, governed b~ the risk of tether

To relieve the maneuvering requirement, one might be prepared to take chances with known debris>10 cm and unguided operational spacecraft, when the risk of an actual collision turns out to be very small.

4.3.2

Avoidanceby operationalSIC

The simpler alternative is to have an uncontrolled (faster) deorbit. The total number of avoidance/alert situations for guided operational spacecraft should be of equal or lesser order of magnitude than in the BAU case. Lifetime is much shorter than for BALI and there is no Future Factor. but the sweep area A~o~p is larger [4.2. I/4.3. I ]. Because of the packing of many avoidance maneuvers in a short time, the annoyance level could be high, but this is something that perhaps can be sorted out, since the objective of the de-orbiting action is eventually beneficial for all parties.

CUt.

A final scenario is the possibiliD of an intentional tether release t~om the remaining s,,stem in case of an extreme emergenc). Some failures that might lead to cut of the tether can be: inadvertent deployment, partial deployment. erroneous upward deployment, excitation ol large oscillations or last minute collision avoidance, in such a case the debris would fall back in the BALI situation, v, hereas the lifetime of the tether, depending on whether the cathode is still attached, is either reduced or increased by an order of magnitude [see 2.2]. Since that would be a second order failure the resulting risk v.ill not affect the comparison. A significant effect might be that release of the tether could mean loss of cxmtrol and theretbre transition from the controlled co_seto the uncontrolled co.se.

4.3.1

Avoidance of operational SIC The height of the sweep area is determined by the sum of tether length L and height SH of a safet), envelop. The x~idth ig determined b) the sum of the tether's oscillation amplitude (half the tether length for a 15 degree libration) and the aspect average of width Sx~ and length St of the safet,,, envelope -- 2"rt*(Sw+StL Assuming Sx~=SH-0.4SI ~se find the following measure-

.4 ,*eep -- ( L ÷SN.)(0. 5~

4.3.3

The risk of a catastrophic event concerns unexpected cut of the tether. If it is assumed that the remains of the tether are automatically disconnected [Figure 2], it can be reasoned that such a cut does not significantly contribute to the comparison with BAU. The risk of tether cut should be small however. A muhi-strand tether with a width of l0 cm is assumed. Catastrophic is any debris of diameter > l0 cm or any non-guided operational satellite. ]'he sweeping width or width of "lethal" interaction is (in simplified form) the diameter of the colliding object. This seems trivial [Figure 3] but is often done wrongly. From data presented in Wile,,' & Wertz 3-" we have determined that the average diameter of objects > l0 cm is about 1.7 m.

Lethal hit left ( ~ extreme

Tether

~ ~! ~ Lethal hit i ~.~# t right extreme I

I

~ - I ~ Width of lethal

D 'interaction

÷2S.)

If one has the opportunity to actively control the tether system's descent, it is likely one also takes responsibility. 1-he total amount of evasive maneuvers (stalling the descent for a single tether length by cutting the current) should not exceed that which is physically possible: one at a time. Nor should it significantly decrease the descent performance, so say less than 25% of the time can be spent on evasive maneuvers. An evasive maneuver for a 10 km tether might take ses.eral days at 800 lan, and about a da) at 400 kin. A rough measure for allowable evasions becomes once every 10 days. Alternatively, if one accepts a significant dela.~ in descent time, one evasive maneuver may occur e v e o --3,4 days.

Break-up collisions

Figure 3: Tether sweeping width

5.

Risk analysis

As an example the method of Section 4 has been applied to a D~ical range of EDTS de-orbits, for a t)13ical 700 kg constellation satellite 3334. We used ETBSim [5.1] and a quick lifetime evaluation tool QrLESim (based on equations ot" Vannarom" e.a. ~'" ). First. an analysis was made assessing the collision risk tbr the BAU and EDTS cases. when avoidance b,wof operational satellites is not taken into account. ]-he a~oidance comparison is made afterwards.

51st L4F Congress

5.1

Delta-Utec Simulation Tool

Below are presented two examples of tethered lifetime reduction conform the suggestion to use a sweep volume for comparison. To use the sweep volume as a measure for debris risk is possible in this case, since a >10 cm debris object is presumed catastrophic for a satellite I~ as well as for a 10 cm wide multi-strand tether. Identical debris densities would have to be multiplied with the sweep volume to come to a risk estimate [Section 4]. Included in the comparison is the concept of disconnecting the tether as soon as a reduced lifetime orbit for the debris has been achieved [3. I. I ].

Analyses have been performed with ETBSim, a 3D/Win95 version of BeadSim 2°. It has been further developed by Delta-Utec to include deployer hardware noise, random environmental disturbances, control algorithms, deployment optimization as well as a conductive tether package. The conductive package exists of: composite tethers (fiat or round conductive/bare/mechanical), electron/ion collection, thermal effects, IRI ionosphere, IGRF (internal), %96 (external) magnetosphere, solar pressure, J2, Sun and Moon gravity effects. Implementation of hardware such as battery charging, hollow cathode and FEA cathode, endmass dynamics and tether torsion is current work.

5.2 Risk of BAU vs. EDTS de.orbit- without avoidance 60

¸ -

-

-7 Skin

m12

50

elecZethef

5 kin_elec tether

I

>,

~4o E

c

._ . . . . .

300

500

700

900 Aitltud• [kin]

1100

1300

Figure 4: Descent rate of 700 kg satelfte - without stability control law, but with current limit.

in Figure 4 the descent rate of a 700 kg satellite, deorbited with an electrodynamic tether, as a function of altitude is given for an equatorial orbit. No control law for stability of the tether at lower altitudes has been applied for this simulation. Nevertheless a current limitation has been accounted for so that the descent rates represent values expected to be achieved when a control law for stability is applied 15. A maximum inclination exists for advantageous use of the tether system (depending on initial altitude and mass of the debris). Maximum obtainable descent rates scale roughly with the cosine of the orbit inclination, or in fact the angle between the system's velocity and the (co-rotating) magnetic field. The relative direction of the magnetic field excites more out-of-plane dynamics at increasing inclination and more intensive control of the current is required for stabilit~ [2.3] at the expense of descent rate. l i m e needed to de-orbit satellites or spent stages is therefbre increasing significantly with inclination, and so will the sweep volume and collision risk.

511

1500

The following assumptions and values have been adopted: • Debris mass: 700 kg~ • Tether length contains an additional 3 km mechanical tether ( I kg); • Maximum lili~time of debris that makes sense for propagation: 150 years [section 41; • Furore Factor: linear increase by factor 4.5 in debris population over the next 150 years [section 4]: • De-coupling of tether+cathode fi'om debris by (non-optimal) swinging release [section 3.21 - tether semi major axis after decoupling decreased by 51. while satellite semi-major axis is increased b3, 0.3• (conductb, e part + mechanical part); i The thrust ofthe tether after cut is conservatively reduced to 20 % ofobtainable thrust [2.21: Electrodynamic tether: eftkctive diameter = 0.8 mm. densit~ = 2 kg/km; • EDITSsubsat: 15 kg: • Solar average: Teso~,hete = 1050 K: • Equatorial orbit; In Figure 5 and Figure 6 the scaled sweep volume (definition see section 4) is plotted versus the altitude for deboost systems with an electrodynamic tether length of respectively 7.5 and 12.5 km. The four lines in these figures represent: • Electrodynamic tether deorbit system (EDlrS) • Decoupled system, sum tether 'and debris: the sum of the sweep x.olume of the debris and of the tetber+subsat after cut. • Deeoupled, debris only: The sweep volume of the debris after decoupling of the system and slight inerease of its apogee due to momentum transfer at cut (- BAt/). • Reduced orbital lifetime of 25 years: lhe sweep volume of a satellite gith a reduced lifetime of 25 )'ears as suggested for EOL operations N. Including Future Factor. In Figure 5 we see that for altitudes above I100 km the sweep volume of the EDTS exceeds that of the satellite when left alone (due to the limit on orbital lifetime [section 4]). This implies that it is better not to accommodate the satellite with a conductive tether of 7.5 krn above II00 kin. In Figure 6 it can be noted that use of a longer tether of 12.5 km shifts this altitude to over 1400 km. Both graphs show, at around 700 km altitude, that sweep volume of the decoupled system becomes smaller than that of the E DTS, suggesting a decoup[ing of the tether and the satellite at this altitude would be advantageous. The

512

51st ~.-IF Coneles~

decoupled satellite at 700 km however is still above the suggested reduced lifetime orbit of 25 years. Releasing the debris below 600 km instantly decreases the remaining sweep volume b}. a factor of 2.5. This would be the optimal solution for an EDTS which is not designed to avoid debris or satelites.

•

•

10

8

E,ec le~er

"

¢

5.3

-

-

=

--

5.3.1 250

500

750

1000

1250

1500

Altlt u d e [kmJ

Figure 5: Sweep volume for 0.8 mm 7.5 km electrodynamic tether (+ 3 km mechanical) without avoidance (deorbit time 1500 Ion ~70 days)

r ~ [--- ~

~

~

J /' ~--~f,~:-o~i~,o, ~]. . . . / - - - - - / .1.,,,~,~,,~,z~o,~., ,=,~ - _ _ _ _ _ . j . . - -

~4

~3 u__.,

~.,,,~_/~_ .,,-~'~

~ _ ~ . ~ " ~

,

2'50

500

~ , , ~ , ~ . . . . a 2. ,~s .

.

.

.

.

.

.

+---

. 750

operational

.

,

~

,

1000

'1250

1500

Albtude [kiln]

Figure 6: Sweep volume for 0.8 mm 12.5 km electrodynamic tether (+ 3 km mechanical) without avoidance (deorbit time 1500 km ,--30 days)

Abo,,e discussion is Ibr an E D I S in an equatorial orbit - its most favorable condition. In case the system is in an inclined orbit the descent rate decreases stronger than the cosine of the inclination. This afli~cts the cur~.es of the E D I S and the ss~,eep volume of the decoupled system. Sweep volume of a 12.5 km EDTS system at i=45 ° resembles that of the 7.5 km tether in equatorial orbit [Figure 5]. Thus an increase in inclination of 45 ° decreases the effective height of the 2.5 km EDTS from 1400 to II 00 km. I h r e e assumptions are intpactmg the results: • 150 Years limitation to orbital lifetime of satellite..,: Fhe exponential trend in the cur~cs ot the dec~upled ~atcllite s~eep ~olume is interrupted b} the assumption that in 15(/}ears dehris ~sill be either removed or ~.asil.,, a~oided.

Avoidingdebris

With about 15000 debris objects in LEO. or a density, of - 1.5 I0 s km 3 and with help of Figure 5 it can be fuund that about 0.013 impact can be expected during the EDTS deorbit from 1500 km, i.e. - I % chance of mission failure. This risk can be either accepted, or one could try, to avoid these objects activel} through current control by telecommand. Then, due to the limited knowledge of the tether position the sweep width is to be increased from 1.7 meters to 5000 m [section 4], leading to 5000/I.7 x 0.013 -40 evasive maneuvers during the --70 days descent. Considering the descent rate of - 2 0 km/day and a total tether length of 10.5 kin. this is (bare b ) possible (some vextra time should be taken into account since the avoidance of one piece of debris should not lead to getting too close to another). 5.3.2

//

Avoiding debt,s and satellites: BAU vs. EDTS

According to the method of Section 4, it has been evaluated how many avoidance maneuvers would be required during the tethered descent, compared to the BAU debris during its Ill, time.

Deoot~p~ed s u m mlh 4

Without such a limitation the EDTS would always be favorable above 700 kin. [he Future Factor: In case lhture space will be more populated (higher Future FactorL the BALI (-decoupled) sweep ,.olume would be significantly higher and the effective range increases. No avoidance of debris and satellites is assumed. This will be the subject of sec'tion 5.3.

Avoiding the safety box of operational satellites

It is assonled that 500 satellites are distributed equally over . ~ , ~ . . • ,.~ ,~-I0 ~ ~ ~r~ - , ," . L LU, I.e, a uenstt~, Ol -. iu Km ". I Re monet O l section 4 predicts the expected number of avoidance maneuvers in Figure 7. It is found that this number is relativel.~ independent of tether length. Again a lower altitude limit is found when comparing to the BALl case, suggesting a decoupling of the tether from the debris at 600 Ion. About I avoidance can be expected per deorbit. This is feasible, although a few points deserve attention: • This ntmlber will increa~ in the future, especially in the case that constellations will become a success. • Around I000 km there is a peak in satellite density, it is not unlikely that several evasive maneuvers will be packed in a small timefi'ame. "

5.3.3

Avoiding the alert box of operational satellites

Some satellites have a large alert box, so it can be argued that such alert boxes need to be avoided as well. With 25 ~,uch satellites, or a density of 2 10 -~=, and an alert box s~idth of 2000 m the results are plotted in Figure 8. It is clear that tethered deorbit ~'stems will be less of a bother to the big important spacecraft than debris under

51st IAF Congress BAU. The number of avoidance maneuvers is low and feasible. (~l~:o~oio lure I ~

a l ~ 46~1s

18 ~6 Av eml4

t

~U

oM an

ce s

1 ---

3~-

08 06 04

0 200

400

600

1000 *,ilir.de

1200

1400

pul

513

Moon) position and solar cycle. Mass-wise they perform better for lower altitude and higher payload mass. Significant mass can be gained for both spent stages and defunct constellation satellites using electrodynamic tethers (EDTS). Figure 9 shows a trend comparison (actual values may be - 3 0 % off). The EDTS mass is based on linearized estimates for the tether system mass (excluding inclination effect) from Forward & Hoyt 34 combined with the work in this paper. The rocket mass is based on Schonenborg 33 who selects an autonomous system for typical constellation satellites. It includes a cluster of AP-HTPB" thrusters, a spin-up thruster and attitude determination and is designed to survive the nominal satellite life (clusters make it failsafe and practical to implement in the satellite design). The result is a propellant mass fraction of 0.66. IO(XI r

--28(X) tcobyrc~V~($PentStaOe)

/I-,=,0.-

]

- - ldOQkO by tnOka

Figure 7: Average number of avoidances of satellites

- - 2800 k,g by EDT$

I ~ko

014 012 m e 01

~m~___

~

B

<004 011"2 ~ J - -

o...Z 2~

A I m , ~ , P=,I 400

600

800

1000

1200

1400

~=tude

Figure 8: Average number of avoidances of alert boxes 6.

Interpretation

of results

What does the research presented here amounts to in terms of comparing mitigation methods? 'Conventional" de-orbit solid rocket boosters for constellation satellites, or extended use of the upper stage burn capability is relatively safe, precise, fast and operationally not very demanding. Tethers can be preferred in cases where they would safe significant mass. Mechanical tethers in LEO score second on the simplicity ranking, since deployment and precise deorbit with very simple technology has been verified and it only lasts several hours. For practical reasons they are hmited to spent stages and they require an interface (and agreement) of the customer. In return, the customer would gain on the average 3% of mass budget. In GTO a spent stage rocket deorbit system requires significantly extended lifetime, which would drive up system mass. The tethered alternative is by mechanical tether, though gain is small and extra operations lasts for 0.5-3 days. Mechanical tethers are safe in lower LEO, but otherwise should only be used in consideration of relative Sun (and

Figure 9: Trendlines for mass comparison (mass tether system = 20 +7_q 10.9 x altitude fro] • mass debris [kg], Isp =300, propellant mass fraction 0PMI;') = 0.66, spent stage PMF=I).

Without controlled deorbit EDTS effectiveness in debris mitigation is less adequate than the rocketed and momentum transfer alternatives. With respect to Business As Usual it is within the same order of magnitude and strongly dependent on altitude earl inclination. It can be improved (and dynamic stability problems avoided) by disconnecting the system around 600 kin. All in all we find a rough application range of 700-1500 km and 0-65 ° inclination. This is compatible with altitude range for the mass advantage with respect to rocketed deboost [Figure 9]. A risk of (large) debris impact on the tether [5.3.1] remains. If however descent rate control will be applied for avoidance, the EDTS system is an attractive alternative concept for debris mitigation. The control of avoiding catalogued debris (> 10 cm) will require more demanding

operational effort.

Fuel: Hydroxyl Terminated Polybutadien, Oxidizer: Ammonium Perchlorate

514

5/st IAF Congress

Results are combined in Table 6 and 7.

LEO

ET Y"

MT Y"

SSO

N

Y"

GTO GEO

N N

Y'. " N

Remark ET (without descent control) 700-1500 km, i<65 °, rocket is more safe, but heaw MT spent stages<500-800 kin. Impact on payload, mass advantage small, Payload should be heavy. ET: Either electromagnetic field line direct=on parallel to velocity (< 2000 kin) or electromagnetic field strength too low (>2000 kin) MT: see above MT: Only under certain conditions. No gravity gradienl/F_M force, no tether deorbit mechanism

Table 6: Application range for tethered debris mitigation. MT=Mechanical, ET=Eleetrodynamic, * can be extended by DUTcther or solar max o r descent control, ** depending on orbit relative to Sun/Moon

spent ~ ' ET MT

17% 41%

Coma~ta" 54% 0%

Dedk:atKI ~ ' 7% ISS/Mir waste disposal

Table 7: Candidates for tethered debris mitigation The results mentioned above have a limited validity because of the assumptions in this paper (150 years debris horizon, LEO increase of factor 3 in I00 years etc.) and because of the following considerations: • A doorbit by EDTS will remove heavy debris at low cost, with good certainty and within a comprcheadible horizon. Removal should be preferred over taking chances on future developments. In fact, future drops in launch cost may very well trigger an exponential rather than linear growth in the space market and occupation. • If tethered doorbit can be delayed until a solar max, it might considerably improve performance. • Today spent stages in LEO arc sometimes deorbited based on an ad hoe decision and fuel leftovers. • For simplicity it has been assumed that tether impact is catastrophic. It should be considered that a tether impact might disable a satellite, but will not be able to disintegrate it, such as a compact piece of debris might do. • There is considerable room for improvement of the current state-of-the-art tether design. At present, currents are limited by deployer technology, plasma contactor capacity and dynamic stability more than by electron collection. The latler can be done very efficiently with light-weight thin and wide bare tether tapes. 7.

Conclusions

Many safe and interesting near-term tether applications exist such as artificial gravity, atmospheric research, sample return capability for Space Station and electrodynamic orbit maintenance of Space Station. However, without a degradable DUtether, because of potential collision risks, there are orbital limitations to the

use of tethers. These limits have been defined in this paper. Within these limits, tethers themselves can be used to de,orbit defunct satellites and spent stages. Mechanical tethers for spent stage momentum transfer are simple, accurate and fast, but require an interface to the payloaql and can provide only limited mass gain. Without descent control electrodynamic tethers of today's design are less effective in terms of debris mitigation and in this sense less adequate than autonomous rocket deboost. One should take into account the tether's large sweeping area, applied to debris collisions and avoidance of safety boxes of operational spacecraft. The advantage with respect to Business As Usual (leaving debris in orbit) is existent, seems marginal, but would increase with progressive assumptions on technology and future use of space. There is little advantage for tethered deorbit of defunct satellites below 600-700 km. For tethered deorbit of higher satellites, it is proposed to disconnect the tether at this altitude. The tether itself will re-enter quickly due to various mechanisms, the satellite will be left in a reduced lifetime orbit (<25 yrs). This will minimize the overall risk. With descent rate control to avoid satellite safety boxes and debris the electrodynamic tether is an attractive solution for debris mitigation. Within the area of application, considerable mass can be saved with respect to rocket deboost. Open development areas are stability control and in-orbit testing of fail-safe multi-strand tethers, both unquestionably required to make electrodynamic tethers attractive alternatives for deorbit. Longer tethers are foreseen than proposed in recent publications. Due to the simplicity and low system cost, early tether deorbit applications may be for smallsats and university satellites. A reason why they might want to deorbit is one that may become a precedent for any future satellite: to avoid liability claims and expensive law suits against a major commercial operator. It is good custom to eliminate pollution after use, not to inhibit a future progressive growth. For some it may become a necessity. Acknowledgments Thanks

to G i u l i a n o

Vannaroni,

Marino

Dobrowolny,

Enrico Lorenzini for the fruitful discussions. Heather & Jenny for their inspiring enthusiasm.

51st L4F Congress

References

515

from

~s Ockels. W.J., Heide, E.J. van der, Kruijff, M., "Space Mail" and Tethers. Sample Return Capabilio' for Space Station Alpha,

Anselmo. L., Rossi, A., Pardini. C., Updated results on the long-term evolution of the space debris environment. Ad~. Space

IAF-95-T.4. I O.' is Bade, A.. Eichler, P., the improved TERESA-concept for th eremoval of large space debris objects, Proceedings of the I ~' European Conference on Space Debris. Darmstadt, Germany,

* Papers by the author can www.delta-utec.com.

be

downloaded

Res., Vol. 23 No. I., pp. 20 I-21 I, 1999. -" Kruijff M.. Heide. E.J. van der. The YES Satellite..4 Tethered Momentum Transfer in the GTO Orbit. Proceedin~ of TTIM 1997. NASA/CP-1998-206900." Forward, R.L. Hoyt, ILP., Failsafe Multiline Hoytether Lifetimes, AIAA-95-2890, July 1995. 4 Barnds. J.. e.a., TIPS: Results of a Tethered Satellite System. Proceedings of TTIM 1997, NASA/CP-1998-206900. • Carroll. J.A., Oldson. J.C., Tethers for Small Satellite Applications. AIAA/USU Small Satellite Conference, 1995. 6 Anselmo, L., Pardini. C., Assessing the impact oforbital debris on space tethers, CNR-CNUCE, 1999. Sabath, D., Paul, K.G. HypervelociO, Impact Erperiments on Tether Materials, TU Munich. 1996. s McBride, N.. Taylor, E.A., The Risk to Satellite Tethers from

1993 z0 Carroll, J.A., Oldson, J.C., SEDS characteristics and capabilities. Proc. 4 ~ int. Conf. on Tethers in Space, pp. 10791090, 1995. -'l Kruijff, M.. Heide, E.J. van der. Tether System Experiment (TSE) Breadboard Test Report, Delta-Utec, September 2000." - Gavira, J., Rozemeijer, H., Muencheberg. S.. The Tether S(stem Experiment, ESA Bulletin 102. May 2000. Beletsky, V.V.. Lenin, E.M., Dynamics qf Space Tether

Systems, 1993. 2~ Dobro~olny. M., Colombo. G., Grossi, M.D., Electrodynamws of Long Tethers in the ,~ear-Earth Environment, Smithsonian Astrophysical Observatory. October 1976. _'5 Sanmartin, J.R., Martinez-Sanchez. M.. Ahedo, E., Bare Wire

Anodes for Electrodynamic Tethers, J of Prop. and Power 9 3,

Meteoroid and Debris Impacts. 2na European Space Debris Conference. 1997. '~ Tether System Experiment Phase .4 Report, ESA/ESTEC. 1999. it, Heide. E.J. van der, Kruijff, M., On the effect of solar pressure on tethers in LEO. Technical Note TNDO201. Delta-Utec, 2000. i pESA Space Debris Mitigation Handbook, Release I.O. 1999 i- Kruijff, M.. Gijsman, P.. Heide, E.J. van der, Opemng the 112/). .for Large. Light and Non-Ha:ardous Space Structures: Report of a search for a Ul'-degradable material, IA F-99-1.3.05.' ~ Chobotov. V.A., Mains, D.L., Tether Satellite System Colltslon Studk,. IAA-98-1AA.6.5,02, 1998.

.~p. 353-360, 1993. Yamagiwa. Y., Takcgahara, H., Nakajima, A.. Quick Descent o~ HII Upper Stage bv Electrodynamic Tether. LAF-99-V.2.05. Alenia Spazio, R~SC Energya, FATS, ESA/ESTEC Contract 11439/95/NL. 1995. :s Kruijff, M, The Young Engineers' Satellite. Flight results and critical analysis of a super-fast hands-on project. IAF-99-P. 1.04.' "~ Bade, A., Eichler, P., The improved TERESA concept for the removal of large space debris objects, ESA SD-0 I. April 1993. 30 Carroll, J.A., Guidebook for .4nalysis of Tether Applications, Martin Marietta Corp., 1985.

'~ Pelaez. J.. Lorenzini, E.C.. Lopez-Rebollal. O.. RuiL M., .t

q GuMehnes and Assessment Procedures ./br Limiting Orbital Debris, NASA Safety Standard 1740.14. ~'~ Wiley, J.L., Wen.z, J.R., Space Mission .4nalysis and Design,

new kind of dk~amw instabili~., in electrodynamw tethers. AAS DO-190, 2000. L~ Ho2~I, R.P.. Forward, R.L., Perlbrmance of the Terminator [ether for Autonomous Deorbit of LEO Spacecraft, AI,,LA-992839. 1999. t, Johnson, L., e.a., Electrodynamic Tethers Ior Spacecraft Propulsion. AIAA 984)983. 17 Oekels, W.J., BiesbroeL R.G.J.. Verduijn. D.F., Enhanced

Efficienck.' of GEO Transfer and Lunar Mtssions by Using the Upperstage Momentum Transferred by Tether, IAF-q6-A.2.01.

2"a Edition 1993. ~ Schonenborg, R.A.C.. Sohd propellant de-orbiting Jor constellation satellites, Delft University of Technology, June 2000. ~4 Forward. R.L., Hoyt, R.P., .4pplication of the Terminator

Tether Electrodynamic Drag Technologv to the Deorbit of Constellation Spacecraft. AIAA-98-349 I, 1998. ~5 Vannaroni, G.. e.a.. Electrodynamic tethers Jbr deorbit applications. IAF Aa,nsterdam. 1999

51~t 14F Congress

516

drag

Air Because of the high area over mass ratio for a tether (~5-10 mZ/kg) drag force is sufficient to get a free tether out of lower LEO in a matter ~ days. Electrodynamic tethers are slightly heavier although they can be designed fiat and thin for good drag performance (-1-5 m'/kg). To largely avoid Space Station in case of failure, lower inclinations are recommended. The solar cycle can affect the density by a factor of 10 to even 100 around 700 km altitude, equivalent to an altitude range for similar drag conditions of ~:50 km at 350 km to ±150 km at 700 kin. Solar pressure Solar pressure, the momentum exchange of the Sun's radiation (1370 W/m 2) and the spacecraft is another effect important for structures with high area over mass ratio, with a difference that it is independent of height and 99.995% norvdissipative: asp- 4.57E-6 x Alto. For circular orbits solar pressure tends to decrease perigee and increase apogee. With rotation of the Sun around the Earth and precession of the orbit itself, the effect will change direction regularly with respect to the semi-major axis. However, in combination with air drag, which increases exponentially with lower altitude, the affect can be ber~fcial for lifetime. The decrease in pedgee height during half a cycle should be sufficient to have the orbit pass through signirmantbj denser areas of the atmosphere. Investigation of this effect has led us to the finding that it is true for orbits with an inclination around 45 ° and altitude up to 800 km TM. For the YES satellitezz (1997), carrying a 35 Ion Dyneema tether on-board, extensive reseemh was performed on the effect of solar pressure in GTO z. Because of the long time spent in apogee, it is very effective in bdnging down pedgea. The launcher, Ariane 5, has a pedgee of 600 kin, too high for a rapid drag induced re-entry. The solar pressure was found to work favorably about 40% of the 24-hour launch window (free flying tether lifetime in GTO of 1-3 months). Dudng the remainder of the day a launch would cause the semi-major axis to be in undesired alignment with the Sun causing a free flying tether to stay in orbit for tens of years. Current in a free flying tether with cathode A free flying tether with its (light-weight) cathode still attached will re-enter quickly. Since no dynamic control is pOSS,hle without a large endmass, the behavior will be unstable. However now the system is without current control, the cuaent collection (and thus the eiectrodynamic drag force) will be maximmed. Ideally, the increase in descent rate with respect to the full system is therefore even larger than the mass ratio of full system over tether+cathode, which may be as large as a factor of 25. This enhanced descent rate will be slowed down and probably, at times, even stopped due to the dynamic instability. In case of a rotating tether, roughly 20% effectiveness would remain, still 5 times faster than the descent the tether was initially designed for. For typicai mission below 1200 Io~, this means a de-orbit time below 10 days. Current in a free flying tether without cathode If for some reason the tether is cut loose from the cathode, the system will be even lighter, typically by a factor 2. However, without a device to emit electrons, a current will still flow, and the circuit is closed by ion collection. Due to the large mass of the ions and low thermal velocity of ions compared to the orbital speed, such collection is much less effective. Nevertheless, for a lignt-weight free flying tether, the effect can be shown to be still significant:

VEMF Electron ............................ c°llecti°ni~ ax Ioncollection J


/ F,o,..=

/+

/

I-le "]'

/

/

1l' Ion/Electron current in free bare tether without cathode The figure shows the principle. Ohmtc resistance can be neglected due to the very low current SimplifyJeg the OML electron collection relationship e.g. from Vannaroni et al.~ and integrating over/~ we find for the maximum current:

1'~"~ : ~n ' ~d ~ g8m~ 1 ~2~e = electron charge [C], no = electron density [m 31, d = tether diameter [m]. b = electromotive voltage per mater [-0.1-0 2 V/m] Doing the same for the tons and setting both currents equal we find: i,//~ = (re,l i n d m - 0.0324. Integretmg the current, and calculating the (maximum) Lorentz drag force via F = I x B L the total maximum drag for the free tether becomes:

F

2_ 475

t '

V m.

200

Gravity gred,ent forces are two order of megmtude larger (~5 cN). they will tend to tend to stretch the (center part of the) tether vertically Also, since there ts no up- or down side. tn a dynamically unstable rotating case. the ton/electron-current will be effective twtce as often as the free cathode case The descent rate ~11 be about 10 times slower than for the nominal mission Thts is equtvelent to a deorbit time of several months Appendix I: Discussion Box I