A LIQUID METAL ION SOURCE FOR SPACE APPLICATION

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A LIQUID METAL ION SOURCE FOR SPACEAPPLICATION

F. Rüedenauer, W. Steiger, H. Arends, M. Fehringer, R. Schmidt

To cite this version:F. Rüedenauer, W. Steiger, H. Arends, M. Fehringer, R. Schmidt. A LIQUID METAL ION SOURCEFOR SPACE APPLICATION. Journal de Physique Colloques, 1988, 49 (C6), pp.C6-161-C6-166.�10.1051/jphyscol:1988627�. �jpa-00228124�

JOURNAL DE PHYSIQUE Colloque C6, supplkment au noll, Tome 49, novembre 1988

A LIQUID METAL ION SOURCE FOR SPACE APPLICATION

F.G. R~~EDENAUER, W. STEIGER, H. ARENDS* , M. FEHRINGER* and R. SCHMIDT'

Austrian Research Center Seibersdorf. A-2444 Seibersdorf. Austria *space Science Department of ESA/ESTEC, NL-2200 Noordwijk, The Netherlands

Abstract - A liquid metal ion emitter, based on indium, for active spacecraft potential control is described. A spacecraft in sunlight charges positively due to photoemission. Under most conditions, the emission of energetic ions into space can reduce the spacecraft potential with respect to the ambient plasma. The ion emitter consists of an array of 5 indirectly heated emitters and a simple focusing system. The typical operating voltage is near 5.5 kV and the nominally emitted current amounts to 10 pA with a beam opening angle of + 300. The source operates at an overal1:secondary power consumption of 0.45 Watts. The design life time of a module is 5000 h. The emitter modules will be flown on the Soviet INTERBALL, the ESA/NASA CLUSTER and the Japanese GEOTAIL mission.

A body embedded in a plasma and exposed to sunlight acquires an electrostatic potential that depends on the photons, surface material and ambient plasma conditions. In the near-earth space a sunlit surface will normally charge to a positive potential. The adverse effects of this charge accumulation on spacecraft are known since long (I/. Under unfavorable conditions in the earth-shadow, the surface potential of a satellite may acquire several kilovolts negative. If some parts of its surface are non-conductive, electrostatic discharges may occur and result in anomalies, malfunction or the total loss of the spacecraft. It is not only such high potentials that are of concern. On scientific spacecraft, usually designed to prevent high potentials, potentials only a few volts above the ambient plasma may disturb the low-energy particle and electric field measurements. Although kfiown and understood since long, no countermeasure has been taken yet. This was mainly due to the lack of small and light-weight ion emitters. Positive. charges must be emitted to neutralize the potential acquired' by a sunlit and conductive surface. The small and reliable liquid metal ion emitters open the possibility to build compact instruments that allow active potential control of spacecraft.

Liquid metal field ion sources (LMIS) have found applications in a number of new technologies, e.g. ion beam lithography, microfabrication and microanalysis /2,3/. In most of these applications the extremely high beam brightness of this source type is used to produce ion .beams with high current densities and small spot diameters (down to a few tens of nanometers /4/). Another intrinsic capability of the LMIS is its high charge efficiency aria its capability to produce ions directly from the condensed phase with low power consumption. This is demonstrated by the development of Caesium field emission ion thrusters /5/. The low power consumption and the small weight of a LMIS makes it an ideal emitter for space application. Laboratory tests on a space-qualified version of an indirectly heated In-LMIS module are in an advanced stage. Low weight, low power consumption LMIS modules are scheduled to be flown on a number of space flight missions.

Article published online by EDP Sciences and available at http://dx.doi.org/10.1051/jphyscol:1988627

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2. Stabilization of the spacecraft electrostatic potential

As mentioned above, the floating potential of a spacecraft is determined by the respective currents leaving and impinging onto its surface. The equilibrium potential Vp is reached with the sum of all currents being zero. Fig. 1 shows typical I-V-curves for a sunlit spacecraft in a near earth orbit: currents due to escaping photoelectrons (Ip) and due to ambient plasma electrons collected by the positive satellite (Ig) determine the floating potential Vg (Fig. 1,left). Contributions from ambient ions as well as secondary emission from the spacecraft can be neglected in tfiis near earth environment. The emission of a constant ion current from the spacecraft adds to the ambient electron current and thus reduces V with respect to the local plasma potential (Fig. 1, right). In adgtion, clamping Vp to the steep part of the I-V-curve also helps stabilizing the potential. In the steep part (right panel of Fig. 1) small current fluctuations A1 hardly affect Vp whereas the same A1 results in a rather big shift in potential AV in the freely floating mode (left panel of Fig. 1). To get even more stable potential conditions the emission of the ion sources on the GEOTAIL and CLUSTER missions will permanently be adjusted according to potential data provided by other onboard experiments.

~ig. 1 Floating potential Vp of a sunlit spacecraft in a near earth orbit with respect to the local plasma potential. left: spacecraft is freely floating; right: a constant positive current is emitted into space Ip: escaping photoelectron current: 1,: ambient plasma electron current

The missions where charge control ion guns are foreseen to be flown require an ion beam of the order of 10 p?k for a total emission lifetime of about 10.000 h. Power and weight limitations dictate that the total secondary power consumption is < 1 W and the weight be less than about 1.5 kg, including electronics. Less than 350 g are left for the source itself. Gallium and Indium LMIS have been under consideration. Indium was chosen because of its relative insensitivity towards oxidation and atmospheric exposure and because of a melting point (1570C) which is above the temperature range which is expected to prevail on a spacecraft in near earth orbit (-10/+40 OC). The melting points of 30% and 280C for Gallium and Caesium respectively yield design problems as they do not allow for collective HV-control. Caesium in addition has a high vapor pressure and its contamination layers are heavily photoemissive, an effect which would counteract charge control.

3. Indirectly heated emitter module

Redundancy considerations require that the total useful lifetime of the charge control unit be divided equally among a number of independent individual ion emitters. Thus, a single module of 5000 h design lifetime at typ. 10 pA emission current contains five individual emitters of 1000 h lifetime each. The emitters are switched on sequentially and emit one at a time until exhaustion of charge material. A schematic view of a five-needle emitter module is shown in Fig. 2. The emitter units are mounted in a structtired slab of extremely low heat conductivity ceramics. The emitter units can be individually heated by resistive heating elements which are connected to ground potential and electrically isolated from the emitter tip potential (indirect heating). A common grounded electrode with five extractor apertures extracts ions from the heated tip. A two-electrode, multi-apertured focusing system forms an external beam of nominally 300 total aperture. The central electrode of this focusing system is at tip potential so that no extra power supply is required. The outer casing of the cylindrical module is at spacecraft ground.

~ i g . 2 Schematic view of the focused indium liquid metal ion source array with 5 indirectly heated emitters

The indivddual emitter units (see Fig. 3) are of the well known "needle in reservoir" type. A tungsten needle with 10 pm tip radius is mounted in a reservoir containing 16 mm3 (0.12 g) of Indium. The tip/reservoir unit is vacuum-brazed to a high-purity alumina tube and hermetically closed at the top end. A miniature resistive heating element with a high positive temperature coefficient of resistivity is bonded into the tube. Thus, the heater temperature can be determined from the heater resistance and

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temperature stabilization of the tip under varying external conditions is possible. The emitter element is 4 mm in diameter and 17 mm in length with a total weight of 9.7 g.

Fig. 3 Individual ion emitter unit with focusing lens

The design of the emitter element itself and the ceramic support slab that accommodates all elements (see Fig. 2) has been optimized for minimum heat consumption. The electrical insulation between the heater and the tip causes a temperature drop between the heater element and its Indium-reservoir but it also drastically simplifies the control electronics. Fig. 4 shows the dependence of heater and tip temperature on heater power. Thermal modeling shows that for a 220°C tip temperature, about 60% of the heat is lost by radiation from the reservoir, - 10% by radiation from the rear end of the heater, - 15% by heat conduction through the heater wires and, finally, 15% by heat conduction through the slab.

The emitters can be exposed to atmosphere for a few days. After exposure they typically start at a firing voltage of 6.5 kV and a heater power of about 0.62 W. A burnt-in source requires a heater power of about 0.44 W. Other tests have proven that such a source can be operated in a pulsed mode with 100% current modulation and 100 ns pulse length /6/. Tab.1 lists the typical steady state operating characteristics.

C- HEATER VOLTAGE X10 1

1 I I I I 8.Wfl 0.1M) 0.208 0.300 0 . W 0.588 0.600

HWLTER POWER (w)

Fig. 4 Tip- and heater temperature in OC vs. heater power (W) for indirectly heated emitters. Heater voltage (Volt x10) is shown in the lower curve.

Tab.1: Operating characteristics of indirectly heated, focused Indium Liquid Metal Ion Source

4. Software for automatic operation

The software allows for completely unattended start-up, operation and shutdown of the ion emitter. It also executes a complete pre-start source checkout, performs error diagnostics and undertakes corrective actions as soon as the source attempts to enter an undesirable operating state. In fact, two simulation programs, respectively called "onboard" and "ground controln, may run in parallel. 9fOnboardw1 executes the autonomous startup, normal operation and shutdown cycle. lUGround control" performs data logging

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and data display functions and allows for operator intervention (change of extraction current, heater power, etc.). The two programs communicate only via two external files and are supposed to simulate the automatic operation of the emitters on a spacecraft and the remote observer on ground. The actively controlled and adjustable source parameters are (a) the drain current and (b) the heater voltage (plus an on/off control capability for tip voltage); the source parameters to be monitored are (a) the emission current, (b) the tip voltage and (c) the heater current.

The programs are written in FORTRAN and implemented on a PDP 11/34 with a CAMAC 1/0 - system.

ACKNOWLEDGEMENTS

F.G.R. and W.S. acknowledge their support by the Austrian Ministry of Science and Research. All authors are grateful for discussions with A.Pederson and H.Svedhem.

REFERENCES

/1/ DeForest, S.E., "Spacecraft Charging at Synchronous OrbitrPP J.Geophys.Res., Vo1.77, 1972, pp. 651-659

/2/ J.Melngailis, J.Vac.Sci.Tech.B, 5 (1987) 469 /3/ F.G.Rudenauer, P.Pollinger, H.Studnicka, H.Gnaser, W.Steiger and

M.J.Higatsberger in "Secondary Ion Mass Spectrometry: SIMS IIIw, A.Benninghoven, J.Giber, J.Lazlo, M.Riede1 and H.W.Werner, Eds.,Springer Verlag (1982), p.43

/4/ R.Levi-Setti, J.Chabala and Y.L.Wang, Scanning Microsc., Suppl.1, 1987. 13

/ 5 / ~.~a;toli, K.v.Rohden and S.P.Thompson, in ltProc. 29. Int. field Emission Symp.lP H.O.Andren and H.Norden, Eds.,(Almqvist & Wiksell, Stockholm 1982), p.363.

/6/ J.Kisse1, H-Zscheeg and F.G.Rudenauer, Appl.Phys. A, in print, 1988