Spacecraft Charging Mitigation Using Mirrors and LEDs

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1 Spacecraft Charging Mitigation Using Mirrors and LEDs Shu T. Lai May 18,

2 OUTLINE Spacecraft charge often to kvs. Why? When? Mitigation using electron emitters. Deficiency! In sunlight, photoemission current exceeds the ambient electron current. Do spacecraft charge in sunlight? Dielectric spacecraft charge to kv in sunlight. Why? Propose using mirrors, VUV lamps, or LEDs, for mitigation. Mirrors and artificial VUV can generate abundant photoelectron current exceeding the ambient electroncurrent. 2

SPACECRAFT CHARGING SPACECRAFT CHARGING: During energetic space plasma events, spacecraft receive excess electrons on the spacecraft surfaces.

3 SPACECRAFT CHARGING SPACECRAFT CHARGING: During energetic space plasma events, spacecraft receive excess electrons on the spacecraft surfaces. DIFFERENTIAL CHARGING Neighboring surfaces of different material properties charge to different potentials. ADVERSE EFFECTS: Charging may disturb the electronic measurements on board, affect the health of the electronic instruments, and terminate the mission. 3

4 HOT ELECTRONS IN THE MAGNETOSPHERE 4

5 5 (Lai and Dell-Rose, JSR, 2001) CAUSE OF SPACECRAFT CHARGING HOT ELECTRONS SATELLITE LANL 97-A

7 MITIGATION METHODS Use electron emitters. Use low-energy ion emitters. Use low-energy plasma emitters. Use spray of water or chemical droplets. Use surface materials of high secondary emission coefficient. Use semi-conducting paint. Each method has advantages and disadvantages! (For review, Lai, S.T., IEEE Trans.Plasma Sci., 31, pp , 2003) 7

8 MITIGATION METHODS Use electron emitters (Often, more efficient emitters are proposed!) Use low-energy ion emitters (Tested in space!) Use low-energy plasma emitters (Tested in space!) Use spray of water or chemical droplets Use surface materials of high secondary emission coefficient Use semi-conducting paint (Tested in space!) Each method has advantages and disadvantages (For review, Lai, S.T., IEEE Trans.Plasma Sci., 31, pp , 2003) 8

9 DEFICIENCY OF THE ELECTRON EMITTER METHOD Differential charging between surfaces is undesirable. Electron emitters emit electrons from the conducting surfaces only. Electrons cannot conduct on dielectric surfaces. Example: Initially, -4kV everywhere, Finally, 0V on conducting surfaces, but -4kV on dielectric surfaces. End Result: Differential charging which is worse than the initial situation. (Lai, 2003) 9

10 DO SPACECRAFT CHARGE IN SUNLIGHT? In sunlight, a conducting spacecraft charges to a few positive volts. A spacecraft with mostly dielectric surfaces can charge to high negative volts in hot space plasmas, even in sunlight. ECLIPSE

11 CAN SPACECRAFT CHARGE TO POSITIVE POTENTIAL? A photon of energy E generates a photoelectron inside a material ( few nm in depth). E forms a distribution with a tail. The Lyα line has 10.2eV. To leave, the photoelectron has to pay a departure tax called Work Function W. W 3 to 5 ev for typical materials. The photoelectron coming out has an energy E ph E W ev. Though photoelectrons are abundant, they are of low energy. They cannot charge a spacecraft to beyond a few low positive volts. If a conducting spacecraft charges to above a few positive volts, most photoelectrons cannot leave. Indeed, conducting spacecraft charge to a few +V in sunlight. SURFACE Photoelectrons

12 SPACECRAFT CHARGING IN SUNLIGHT The ambient electron flux J e at geosynchronous altitudes is 0.115x10-9 A/cm 2 [Purvis and Garrett, 1984]. The outgoing photoemission flux J ph from typical surfaces is 2x10-9 A/cm 2 [Stannard, et al., 1981]. Outgoing photoelectron flux J ph Incoming electron flux J e X 20. How can spacecraft charging (to negative volts) occur in sunlight? J ph J e SURFACE 12

13 CAN SPACECRAFT CHARGE TO NEGATIVE VOLTS IN SUNLIGHT? Kirchhoff s law states that current balance holds at every circuit junction. The spacecraft potential φ is determined by current balance: I ( f )- I ( f )- I ( f ) = 0 e i ph where I e (φ) is the net incoming electron current, I i (φ) the ion current, and I ph (φ) the outgoing photoelectron current. If I e (φ) << [I i (φ) + I ph (φ)], there is no solution φ. So, it seems that charging to negative volts in sunlight is impossible! However, spacecraft do charge in sunlight to negative volts!! 13

14 Mott-Smith and Langmuir Model of Current Collection At geosynchronous altitudes, it is often a good approximation to use the Mott-Smith and Langmuir [1926] formula for describing charge collection in the orbit-limited regime. The current balance equation can be written as follows. where a e e e i ( i i) ph I (0)[1 - d+ h ]exp(- q f / kt )- I (0) m1 - q f / kt - I ( f ) = 0 and δ 0 + η = 0 dee f ( E) δ( E) + η( E) dee f ( E) qiφ Ii( φ ) = Ii(0) µ 1 kti Here, q e and q i are the electron and ion charges, k Boltzmann s constant, f(e) the distribution of the ambient electrons, T e and T i the ambient electron and ion temperatures respectively, and I ph (φ) is the photoemission current from the spacecraft at potential φ(<0). μ is a factor related to dimension (μ = 1 for a sphere, μ = 1.1 for an infinite cylinder, and μ = 1 for a plane [Mott-Smith and Langmuir, 1926; Laframboise and Parker, 1973; Lai, 1994]), α = 1 for a sphere, ½ for a cylinder, and 0 for a plane. α 14

15 Calculated Spacecraft Potential for Various Photoemission Currents SPACECRAFT POTENTIAL (-kv) I / I (0) = 0.4 ph e I / I (0) = 0.5 ph e 1-D 2-D 3-D ELECTRON TEMPERATURE (kev)

16 16 (Lai and Dell-Rose, JSR, 2001) CAUSE OF SPACECRAFT CHARGING HOT ELECTRONS SATELLITE LANL 97-A

17 POTENTIAL CONTOURS MONOPOLE-DIPOLE MODEL SUN (Besse and Rubin, 1980) 17

SATELLITE CHARGING IN SUNLIGHT MONOPOLE-DIPOLE POTENTIAL CONTOURS SUN DIRECTION SATELLITE For a satellite with dielectric surfaces, the dark side can charge to high negative kv.

18 SATELLITE CHARGING IN SUNLIGHT MONOPOLE-DIPOLE POTENTIAL CONTOURS SUN DIRECTION SATELLITE For a satellite with dielectric surfaces, the dark side can charge to high negative kv. Potentials contours wrap to the sunlit side with barrier formation. - kv Satellite Radius R = 1 Monopole Strength K = 1 Dipole Strength A = 0.6 Barrier Location = 1.2 Barrier Potential = 16.7V (Lai, 2003) 18

19 POTENTIAL SADDLE POINT AND BARRIER HEIGHT The monopole-dipole potential 1 Acosθ φθ (, r) = K r r 2 The potential barrier is located at (0 o,r S ) which gives R S =2A Let us denote R as the satellite radius. Since the saddle point can not be within the satellite, R S >R, and 2A>R. The potential barrier height is B o dφ(0, r) dr o o B= φ(0, R ) φ(0, R) = K Photoelectrons and secondary electrons have a few ev in energy. The barrier height B can be small (a few V) to block the photoelectrons and secondary electrons. i.e. (2A-R) can be small. i.e. A R/2. S r= R S = 0 ( R 2A) 2 4AR 2 Shu T. Lai

20 RATIO OF POTENTIALS The ratio of front to back potentials is 1 A o K 2 φ (0, R) R R R A Ratio = = = o φ(180, R) 1 A R+ A K + 2 R R Since the potential barrier B can be small compared with the potential at the back. Substituting A R/2 into the above equation, we obtain Ratio o φ (0, R) R A R R/ 2 1 = = o φ(180, R) R+ A R+ R/2 3 Thus, the Ratio between the potentials on the sunlit side and the dark side is simply 1/3 approximately. Shu T. Lai

21 OBSERVED SPACECRAFT POTENTIAL AND ELECTRON TEMPERATURE ECLIPSE OBSERVATION: The ratio of the potential curve in eclipse to that in sunlight is about 1/3. SUNLIGHT (Lai and Tautz, 2006) Shu T. Lai

22 SATELLITE March 13 28, OBSERVATION: The ratio of the potential curve in eclipse to that in sunlight is about 1/3. (Lai and Tautz, 2006) Shu T. Lai

23 PAUSE FOR A SUMMARY Spacecraft charging is caused by hot (kevs) electrons. There are mitigation methods; they have advantages and disadvantages. From time to time, more efficient electron emitters are proposed. Photoelectron flux exceeds the ambient electron flux. Photoelectrons are of low energy (a few ev). In sunlight, a conducting satellite charges to a few positive V only. In sunlight, a satellite with mostly dielectric surfaces forms potential barrier on the sunlit side. In sunlight, a satellite with mostly dielectric surfaces can charge to high V. 23

24 NOVEL MITIGATION METHOD USING MIRRORS IDEA: If one can mitigate the charging on the dark side, the potential wrapping disappears. Without potential wrapping, the barrier on the sunlit side disappears. Without the barrier, photoelectrons can leave. If the photoelectrons can leave, the charging of the entire satellite is mitigated. 24

25 MIRROR REFLECTANCE The photoemission current generated by typical surfaces in sunlight is 20 times larger than the energetic electron current. Aluminum mirror with reflectance of 86% [Keski-Kuha et al., 1999] is used on Hubble for the telescope, not for mitigation. There is no blockage for the photoelectrons generated from the negatively charged surfaces on the satellite s dark side. With 86% sunlight reflected from the mirrors onto the dark surfaces, the photoelectron current generated exceeds that of the ambient electrons. Reflectance of aluminum surface as a function of incoming photon energy (plotted by using data taken from CRC Handbook of Physics and Chemistry, CRC Press, 2003.) 25

26 CHARGING OF MIRRORS The photoelectron flux emitted from a surface is given by J ( ph ωα, ) = J ( 0 ω ) Y ( ωα, ) [ 1 R ( ωα, ) ] where J 0 is the incident light intensity at frequency ωω and incidence angle αα, Y(ω, α) is the photoelectron yield per absorbed photon, and R(ωω, αα) is the reflectance. If R 1, the photoelectron flux J ph must be small. CONJECTURE: A mirror should charge in sunlight as if it were in eclipse! (Lai, JGR, 110, pp , 2005) 26

27 MITIGATION OF MIRROR CHARGING Two simple methods: 1. Use sharp spikes at the back of the metallic mirrors. Silos can be used to reduce ion sputtering. e - e - e - SURFACE 1. Attach a non-reflecting metal foil to increase the sunlit area for more photoemission. e - R << 1 MIRROR e - R << 1 27

28 VUV PHOTONS FROM LAMPS GENERATE PHOTOEMISSION Concept of using Vacuum Ultraviolet (VUV) Photons to remove surface electrons on the satellite s dark side. 28

29 PHOTOELECTRONS EMITTED FROM CONDUCTORS AND DIELECTRICS 29

30 ARTIFICIAL VUV LAMP AND VUV LED For photoemission, the photon needs an energy above the work function W plus some attenuation loss. W 4 ev for most surface materials in space. The photon wavelength has to be shorter than about 300 nm. Commercial VUV lamps and LED are available (e.g., Hamamatsu, Nikkiso, of Japan, Osram, USA). Using the Osram xenon lamp (170nm), Dickson et al [2009] reported that the measured photoemission flux J ph from a Pt surface was 0.01A/m 2. The J ph exceeds the ambient electron flux J e (= 2x10-5 A/m2) at geosynchronous altitudes. 30

32 VUV LAMPS AVAILABLE COMMERCIALLY Measured spectra of photon intensity (in relative units) of OSRAM Xeradex [left] and Hamamatsu H2D2 [right] lamps. The cover glass made with MgF 2 is for improved transmittance. 32

33 RATE OF PHOTON GENERATION FROM A VUV LAMP Using the photon spectrum f L (E), the photoelectron flux J ph is given by J = de f 0 L( E) Y ( E) where Y(E) is the photoelectron yield at the photon energy E. Absolute units on the y-axis are lacking, can we do an estimate of the photoelectron current? Take the Osram f L (E) (left) spectral line at 170nm (VUV, 7.7eV). Osram lamp output power P = 40 Watt. The number N ph of VUV photons per sec is obtained by equating the energy rates: 7.7 N ph which gives N ph = 3.2x10 19 photons.s -1. = P 33

34 THE PHOTOELECTRON PRODUCTION RATE We have obtained a photon production rate of N ph = 3.2x10 19 photons.s -1 The yield function Y(E) at 7.7 ev is about 10-3 photoelectrons per photon (Grard, 1972). The rate of photoelectron production rate is given by N ph Y(E) = 3.2x10 16 e - s -1. which gives a photoelectron current J ph = A. If the photoelectrons are spread out over 1 m 2, the current density is near Dickson s measured result: 0.01 A m (Feuerbacher and Fitton, 1972)

35 REFERENCES Besse, A. and A. Rubin, A simple analysis of spacecraft charging involving blocked photoelectron currents, J. Geophys. Res., Vol.85, pp , (1980). Hastings, D. and H. B. Garrett, Spacecraft Environmental Interactions, Cambridge University Press, Cambridge, UK, (1996). Lai, S.T. and D. Della-Rose, Spacecraft charging at geosynchronous altitudes; New evidence of the existence of critical temperature, J. Spacecraft & Rockets, Vol.38, No.6, , (2001). Lai, S.T., A Critical overview on spacecraft charging mitigation methods, IEEE Trans. Plasma Sci., Vol.31, No.6, , (2003). Lai, S.T., Charging of mirrors in space, J. Geophys. Res.- Space Phys., Vol.110, A01, , (2005). Lai, S.T. and M. Tautz, Aspects of spacecraft charging in sunlight, IEEE Trans. Plasma Sci., Vol.34, pp , (2006). Lai, S.T., Fundamentals of Spacecraft Charging, Princeton University Press, Princeton, NJ, (2011a). Lai, S.T., et al, Spacecraft Charging, Progress in Aeronautics and Astronautics Series, S.T. Lai (ed), AIAA Press, (2011b). Lai, S.T. and K. Cahoy, Trapped photoelectrons during spacecraft charging in sunlight. IEEE Trans. Plasma Sci., 43(9), pp , (2015). Lai, S.T. and K. Cahoy, Spacecraft Charging, in Encyclopedia of Plasma Technology, pp , CRC Press, (2017). Feuerbacher B. and B. Fitton, Experimental investigation of photoemission from satellite surface materials, J. Appl. Phys., vol. 43, no. 4, pp , (1972). 35

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Project report. Spacecraft Charging and Mitigation Methods. By Muhammad Azam. Abstract

Project report. Spacecraft Charging and Mitigation Methods. By Muhammad Azam. Abstract Umeå University October 7, 2009 Department of physics Space physics 7.5 ECTS. Project report Spacecraft Charging and Mitigation Methods By Muhammad Azam Muhammad.azam1@yahoo.com Supervisor Kjell Rönnmark