Space Environment and Satellite Systems
Understanding Spacecraft Electrical Anomalies: Theory and Experiments Characterizing Hypervelocity Impact Plasma Dynamics Nicolas Lee June 8, 2012
Spacecraft Threats
1
Impact-Related Anomalies Olympus
Landsat 5
ADEOS II
ALOS
JASON-1
1993
2009
2003
2011
2005
ESA
NASA
JAXA
JAXA
NASA
Numerous others with unexplained electrical failures: Galaxy 15, Fengyun-1, AusSat-A3, Intelsat 511, Telecom 1B, Intelsat 510, Arabsat 1-A, Anik-D2. . . NRC report (2011) recommends studying “. . . effects of plasma during impacts, including impacts of very small but high-velocity particles.”
2
Impact Plasma Formation
Impact
Plasma formation
Initial expansion
RF Emission
The phenomenon of electrical damage from hypervelocity impact (of meteoroids or orbital debris) on spacecraft is not well understood Impact energies are great enough to produce a plasma: 3.48 Charge Q ∝ m1.02 proj × vimpact
3
Measurements and Models
Impact
Plasma formation
Initial expansion
RF Emission
Time
Model #1 Model #2
Measurement
4
Contributions In order to enable characterization of the risk of electrical damage to spacecraft from hypervelocity impact, I made the following contributions: Developed physics-based models of plasma expansion to capture relevant and previously ignored dynamics Designed and built novel plasma sensors and conducted ground-based experiments to measure impact plasma expansion Produced first measurement of impact plasma temperature accounting for internal electrostatic forces and of plasma composition from impacts on spacecraft surfaces
5
Outline Background Theory and models Impact experiments Plasma measurements Conclusions
6
Outline Background Impactors — meteoroids and orbital debris Space environmental conditions Previous work — impact studies Theory and models Impact experiments Plasma measurements Conclusions
7
Hypervelocity Impactors Meteoroids: Smaller than 0.3 m diameter (including dust) Shower (associated with parent body) and sporadic (background) sources Parents are cometary (icy, 50–70 km/s) or asteroidal (rocky, 20–30 km/s) Orbital debris: Human-made objects primarily in Low Earth Orbit (LEO) Range from paint chips to rocket bodies Slower than meteoroids (7–11 km/s) but comparable momentum
8
Spacecraft Charging Plasma bombardment vs. photoemission Low Earth Orbit (LEO): +10 V to −700 V Geosynchronous Earth Orbit (GEO): +30 V to −28000 V
9
Ground-Based Accelerators Ground-based accelerators unable to achieve speeds and masses representative of meteoroid and debris populations Test chamber size and pressure limit fidelity of environmental conditions
10
Previous Hypervelocity Impact Research Friichtenicht and Slattery (1960s) — Accelerator development, impact ionization Eichhorn (1970s), Grun (1980s–1990s), et al. — Impact flash, sensor calibration Burchell, Ratcliff (1990s) — Plasma yield and energetics
Ratcliff et al. [1997] 11
Relevant Plasma Parameters Thermal speed: Composition and temperature s vth,j = Expansion rate
kB Tj mj
Ion acoustic speed: Density drop off
r vs =
Plasma frequency
γekB Te + γikB Ti mi
Plasma frequency: s
Radiated spectrum
ωp,e =
ne e 2 0me
12
Plasma Temperatures What is temperature? Average energy of thermal motion (vs. bulk flow) Units - measured in electron volts (1 eV = 11600 K) Issues: choice of population, distribution, species, direction 0.35
f(v) [1/(km/s)]
0.3
10 eV 5 eV
0.25 0.2 0.15 0.1 0.05 0 −10
−5
0 Speed [km/s]
5
10
13
Outline Background Theory and models Plasma dynamics Single particle motion Numerical 1D simulation Impact experiments Plasma measurements Conclusions
14
Plasma Equations of Motion Fluid momentum equation: ← → ∂~v ~ ~ ~ ~ mn + ~v · ∇ ~v = qn E + ~v × B + mn~g − ∇ · P + mnν~v ∂t Collisions Gravity Lorentz Inertial Pressure & Viscosity Particle approach: ~ int + E ~ ext F~ = m~a = q E Geomagnetic field of ∼ 48 µT — negligible for ions, electrons at the speeds and distances of interest Internal electric field dominates when plasma is dense (Debye shielding) External electric field dominates when plasma expands and rarifies
15
Plasma Scale Lengths Ion Larmor radius RL,i
Scale length [m]
101 1m
Electron Larmor radius RL,e
10-1
RPA position dRPA
10-2
Internal electric field Eint ≈ Eext
1 mm 10-4 -5
10
1 µm 10-7
Debye length λD ≈ r Collision frequency νe,i ≈ ωp,e Crater radius Particle radius
10-8 1 nm 10-10
Atomic radius 16
1D Single Particle Motion
F = m¨ x = qEext
~ ext E
Sigrid Close 2012
Position
Current
Assume plasma diffuse enough that each ion/electron moves independently Integrable to get closed form solutions for time-of-flight
Time
17
1D Single Particle Motion Initial speed required to reach distance d in time t:
Current
d qj Eextt v0(t) = − t 2mj
Sigrid Close 2012
~ ext E
Position
t
d v0 Time
18
Initial Speed Distribution 10 eV distributions:
Maxellian or half-Maxwellian speed distribution:
14 Fe W
fq (v) = ηj Q
2mj mj v 2 exp − πkB Tj 2kB Tj
where
fq(v) [fC/(km/s)]
12
s
10 8 6 4 2
Z
v
r fq (u)du = ηj Q erf
0
mj v 2kB Tj
0
0
2 4 6 8 Expansion speed [km/s]
10
Plasma current at distance d, assuming Maxwellian distribution: r X η j Q r mj mj I(t) ≈ erf v0(t) − erf v0(t + δt) δt 2kB Tj 2kB Tj j
19
Initial Speed Distribution Speed distribution causes plasma to disperse
Sigrid Close 2012
Position
Current
Higher temperature results in wider current pulse
Time
20
Numerical 1D Simulation
F = m¨ x = q (Eint + Eext) Assume 1D motion of plasma discretized into shells of spherical caps Each shell imposes an internal electric field on other shells Integrate numerically to get space-time trajectories
21
1D Conical Geometry Define cone based on plume angle and initial plasma radius Initial sphere becomes keystone-shaped plug
22
Model Strategy Assume initial conditions
Propagate plasma dynamics
Find charge density in space and time
Determine plasma current at some distance
Generate synthetic measurement
23
Applying Models
Impact
Plasma formation
Initial expansion
RF Emission
Time
SPM Numerical Single Particle Motion: Quick iteration
Measurement
Solve for composition and measured temperature distribution 24
Applying Models
Impact
Plasma formation
Initial expansion
RF Emission
Time
SPM Numerical Numerical 1D Simulation: Longer runtime
Measurement
Detailed study of certain configurations 25
Outline Background Theory and models Impact experiments Accelerator facility, chamber, targets Plasma sensor designs Plasma measurements Conclusions
26
Ground-Based Hypervelocity Impacts
PSU 2 MV Terminal Detector QP Dust Source
Detector 1 & 2 Chamber
Particle Deflector
Belt
Potential Rings
Positively charged iron dust particles accelerated to 1–70 km/s Impact observed by plasma, optical, and RF sensors
27
Experimental Configuration 1.4 m diameter vacuum chamber Operating pressure ≤ 10-5 mbar Optical sensor Plasma sensors
Particle beam line Target
Patch antennas
28
Impact Targets
Tungsten — baseline target Solar cells Optical solar reflectors Solar panel substrate Electrical bias applied to targets to simulate spacecraft charging
29
Plasma Sensors Plasma particles impinge on metal collecting surface generating net current Grids select for particle species and energies
Repeller grid Threshold grid Suppressor grid Vout+ VoutTransimpedance Differential amplifier driver
Housing ground
30
Amplifier Design Current converted into voltage signal in modular amplifier board Frequency response includes 0.5 ns cable propagation delay
Gain [mV/nA]
(4 × 1014)s mV GRPA(s) = (s + 16.8 × 106)(s + 1.78 × 106)(s + 50) nA 2
10
0
10
Model Measured
−2
10
4
Phase [deg]
10
5
10
6
10
7
10
0 −180 −360 4
10
5
6
10 10 Frequency [Hz]
7
10
31
RPA Calibration
RPA calibration at Lockheed Martin Advanced Technology Center Electron gun built to generate current source 32
RPA Experimental Geometry
RPA 1
RPA 2
RPA 1 at 85 mm from impact point RPA 2 at 65 mm from impact point Similar angle on opposite sides of beamline
33
Faraday Plate Array (theta)
Uses RPA transimpedance amplifier Copper collecting plate without grids Spanned four (actually three) angular positions at constant range
34
Faraday Plate Array (range)
Approximately constant angle to target normal Approximately constant solid angle Spanned four range positions
35
Impact Measurements
SRI [V]
80
50 0 −50 −100 −150
0 −0.5
−20
0
20
−20
0
20
1 PMT [V]
0.5
PA916 [mV]
60
RPA [mV]
1
PA315 [mV]
QP [mV]
100
0.5 0 −20
0 Time [us]
20
−20
0
20
−20
0
20
5 0 −5
10 0 −10 −20
0 Time [us]
20
vproj = 39.4 km/s; mproj = 1.45 fg; Target = Active tungsten; Bias = +1000 V
36
Models and Measurement
Impact
Plasma formation
Initial expansion
RF Emission
Time
SPM Numerical
Measurement
37
Outline Background Theory and models Impact experiments Plasma measurements Plasma composition and temperature Impacts on positive and negative targets Impacts on metal and spacecraft targets Conclusions
38
Composition: Positive Tungsten Multiple positive peaks Plasma plume focused into one RPA H C O Na Fe W 600 RPA 1 RPA 2 RPA [mV]
400
200
0
−200
0
2
4 6 Time [us]
8
10
vproj = 36.3 km/s; mproj = 3.19 fg; Target = Active tungsten; Bias = +1000 V
39
Composition vs. Impact Speed
RPA [mV]
1500
1000
500
0 40
20 Impact speed [km/s]
0
1
2
3
4
5
6
7
Time [us]
40
Composition: Positive Solar Cell Fe and SiO2 with either SiO or K No velocity-based transition in composition
41
Negative Target Bias Fast bipolar response Sometimes a second peak microseconds later 400 RPA 1 RPA 2
RPA [mV]
200 0 −200 −400 −600 −2
0
2
4 6 Time [us]
8
10
vproj = 13.5 km/s; mproj = 26.5 fg; Target = Tungsten; Bias = −1000 V
42
Secondary Electron Emission High-energy electrons impinging on metal surface can bounce off and/or knock off another electron Secondary electron emission yield δ ≡ Iout/Iin Using data from Dekker (1958) for SEE from electron bombardment on tin: Secondary electron yield δ
1.4 1.2 1 0.8 0.6 0.4 0
500 1000 Primary electron energy [eV]
1500
43
Negative Ion Presence
50
FPA [mV]
0 −50 −100 90 mm 60 mm 50 mm 30 mm
−150 −200
0
2
4 Time [us]
6
8
vproj = 4.55 km/s; mproj = 5.62 pg; Target = Tungsten; Bias = −300 V
Second peaks staggered in time based on distance from impact Timing indicative of oxygen ion
44
Spacecraft Surfaces: Negative Bias
Solar Panel
Solar Cell (uncoated) 50
40 RPA 1 RPA 2
0
−100 −150 −200
0
5
10 Time [us]
15
20
−250
100
−20 −40
0 −100
−60
−100
−150
40 RPA 1 RPA 2
−200
−80 0
5
10 Time [us]
15
20
−100
0
5
10 Time [us]
Target Solar Panel Solar Cell (uncoated) Solar Cell (conductive) OSR (standard) OSR (conductive)
15
20
−300
RPA 1 RPA 2
20
0
−50
RPA [mV]
RPA [mV]
RPA [mV]
−50
200 RPA 1 RPA 2
20
0
OSR (conductive) RPA [mV]
RPA 1 RPA 2
OSR (standard) RPA [mV]
50
Solar Cell (conductive)
0 −20 −40 −60
0
5
10 Time [us]
15
20
−80
0
5
10 Time [us]
15
20
Negative Ions No Yes No Two species Yes
45
Temperature Dependence of Plasma Signal Meas. 100 eV 50 eV 20 eV 5 eV 1 eV
RPA output [mV]
800 600 400 200 0 −200 −400 0
2
4
6
Time [us]
46
Temperature Dependence of Plasma Signal
RPA output [mV]
800
Meas. Best fit
600 400 200 0 −200 −400 0
2
4
6
Time [us]
Best fit:
Composition [%] Temperature [eV]
H 25.4 16.3
C 11.9 8.4
O 9.0 3.3
Na 2.9 1.2
Fe 31.0 17.9
W 19.8 22.0
47
Temperature vs. Impact Speed
35
H C O Na K Fe W
Temperature [eV]
30 25 20 15 10 5 0 0
10 20 30 Impact speed [km/s]
40
Temperatures of bulk constituents ∼10 eV Significantly lower temperatures in surface contaminants
48
Plasma Scale Lengths Ion Larmor radius RL,i
Scale length [m]
101 1m
Electron Larmor radius RL,e
10-1
RPA position dRPA
10-2
Internal electric field Eint ≈ Eext
1 mm 10-4 -5
10
1 µm 10-7
Debye length λD ≈ r Collision frequency νe,i ≈ ωp,e Crater radius Particle radius
10-8 1 nm 10-10
Atomic radius 49
Numerical Simulations T ∼ 20 eV 0.08
Position [m]
0.06 e− H C O Na Fe W
0.04
0.02
0 0
1
2 3 Time [us]
4
5
Hotter plasma results in too much dispersion Ion initial temperature at least an order of magnitude lower than final temperature
50
Summary of New Results 1. Initial temperature is at least an order of magnitude colder than previously reported → Low temperature indicates slower initial plasma expansion allowing for more RF emission than previously expected 2. Composition of impact plasmas from glass spacecraft targets has low dependence on impact speed → Plasma generation mechanism may be different than for metallic targets 3. Negative ion formation has strong dependence on target material → Certain spacecraft components may be more prone to impact-related anomalies
51
Contributions In order to enable characterization of the risk of electrical damage to spacecraft from hypervelocity impact, I made the following contributions: Developed physics-based models of plasma expansion to capture relevant and previously ignored dynamics Designed and built novel plasma sensors and conducted ground-based experiments to measure impact plasma expansion Produced first measurement of impact plasma temperature accounting for internal electrostatic forces and of plasma composition from impacts on spacecraft surfaces Publications Lee, N. and S. Close (2012), Understanding spacecraft failures by characterizing hypervelocity impact plasmas, URSI-NRSM, Boulder, CO. Lee, N. et al. (2012), Measurements of freely-expanding plasma from hypervelocity impacts, International Journal of Impact Engineering, 44, 40–49.
Lee, N. et al. (2011), Study of hypervelocity impact plasma expansion, in 3rd AIAA Atmospheric Space Environments Conference, AIAA, Honolulu, HI. Close, S., P. Colestock, L. Cox, M. Kelley, and N. Lee (2010), Electromagnetic pulses generated by meteoroid impacts on spacecraft, Journal of Geophysical Research, 115 (A12328). 52
Outline Background Theory and models Impact experiments Plasma measurements Conclusions Implications of thesis contributions Future work
53
Implications for Spacecraft Existing spacecraft: modify operation to increase safety Negative target prone to discharge-like effect Ensure impact surface is also in sunlight Future spacecraft: design for reduced risk Apply active discharging techniques at critical locations prone to high electric fields Optimize spacecraft geometry (e.g. payload placement) to design for favorable charge condition on likely impact surfaces
54
Future Work Modeling and simulation CFD/PIC simulations to go deeper in plasma formation and initial expansion Higher-fidelity 3D expansion models Ground-based experiments Measure actual effect of impacts on circuits Improved plasma sensors Active plasma modification Light gas guns
55
Future Work: In Situ Measurements MEDUSSA and MORGANA CubeSats Meteoroid, Energetics, and Debris Understanding for Space Situational Awareness Meteoroid On-orbit Research in Geospace for Advancing Near-Earth Awareness
Equivalent sensor suite to measure real meteoroid impacts
Pumpkin, Inc. 56
Zero-Gravity Testing
57
Weatherproof Spacecraft
58
Far Future: Weatherproof Starships
59
Slide Index 1 2 3 4 5 8 9 10 11 12 13 15 16 17 18 19 20 21 22 23 24 25 27 28 29
Spacecraft Threats Impact-Related Anomalies Impact Plasma Formation Measurements and Models Contributions Hypervelocity Impactors Spacecraft Charging Ground-Based Accelerators Previous Hypervelocity Impact Research Relevant Plasma Parameters Plasma Temperatures Plasma Equations of Motion Plasma Scale Lengths 1D Single Particle Motion 1D Single Particle Motion Initial Speed Distribution Initial Speed Distribution Numerical 1D Simulation 1D Conical Geometry Model Strategy Applying Models Applying Models Ground-Based Hypervelocity Impacts Experimental Configuration Impact Targets
30 31 32 33 34 35 36 37 39 40 41 42 43 44 45 46 47 48 49 50 51 52 54 55 56
Plasma Sensors Amplifier Design RPA Calibration RPA Experimental Geometry Faraday Plate Array (theta) Faraday Plate Array (range) Impact Measurements Models and Measurement Composition: Positive Tungsten Composition vs. Impact Speed Composition: Positive Solar Cell Negative Target Bias Secondary Electron Emission Negative Ion Presence Spacecraft Surfaces: Negative Bias Temperature Dependence of Plasma Signal Temperature Dependence of Plasma Signal Temperature vs. Impact Speed Plasma Scale Lengths Numerical Simulations Summary of New Results Contributions Implications for Spacecraft Future Work Future Work: In Situ Measurements
60
57 58 59 60
Zero-Gravity Testing Weatherproof Spacecraft Far Future: Weatherproof Starships Slide Index
61
