Spacecraft bus voltage and power
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Spacecraft bus voltage and power 0.1 1 10 10 2 10 3 10 4 10 100 1000 DSCS II RCA SATCOM MILSTAR TDRS SKYLAB SPACE TELESCOPE LEASAT ISS HS702 space factory 、 space hotel Spacecraft power (kW) Bus voltage(V) EOS-AM ETS-8 Before mid-90s After mid-90s After 2010 experimental solar power satellite severe plasma interaction 400V is necessary for next-generatio 1MW large space platform after ISS
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experimental solar power satellite. 4. 10. 3. 10. space factory、 space hotel. ISS. 2. 10. Spacecraft power (kW). HS702. SKYLAB. 10. MILSTAR. SPACE. ETS-8. TELESCOPE. TDRS. Before mid-90s. 1. EOS-AM. LEASAT. RCA SATCOM. After mid-90s. DSCS II. After 2010. 0.1. 10. 100. - PowerPoint PPT Presentation
Transcript of Spacecraft bus voltage and power
Spacecraft bus voltage and power
0.1
1
10
102
103
104
10 100 1000
DSCS II
RCA SATCOM
MILSTAR
TDRS
SKYLAB
SPACETELESCOPE
LEASAT
ISS
HS702
space factory 、space hotel
Spa
cecr
aft p
ower
(kW
)
Bus voltage(V)
EOS-AM
ETS-8
Before mid-90sAfter mid-90sAfter 2010
experimentalsolar power satellite
severe plasma interaction
400V is necessaryfor next-generation1MW large space platform after ISS
Spacecraft potential in LEO
load
φ1
φ2
φ1
electrons
ions
Varray=φ1 +φ2
φ2 ≈−Varrayd
plasmapotential+ +++ + +
+
+
+
+++++++
++
+
+
+
Charge accumulated over the total area
φ2 =φs
φinsl≈−κTe
Q=εdVarrayA
d+12CcgVarray
φcg≈−κTe
ΔV=φinsl−φs ≈−κTe +Varray≈Varray
ΔV
+
+
+
+
+++++
How an arc occur?
e
ii
ii e
ee
sheathcoverglass (dielectric)
solar cell
interconnector (conductor)substrate(dielectric)
triple junction
e
+ + + + +
eeee
ee
+++
+
+ + + + + + + + + + + +
e
+ + + + + + + + + + + ++++ +
++
eee
ee
e
+
+
+
+
+
+ + + +
+
++
+++
+ ++
Charging of coverglass by ions
Field intensification at triple junction
Field emission of electrons
Ionization of desorbed gas
Neutralization of coverglass charge
Repeated arcing and charging process
Sustained arc
- + - + - +
triggerarc
growth of arc plasma
sustainedarc formation
(a) (b) (c)
Test facility
R
B
G
7cm
3.5cm
21cm
P bus bar
N bus bar
inter-connector
21cm
18cm
RTV
Xe ionthruster
Large-scaleDeployableReflector(LDR)
7.3m
40m37m
2.35m
2.45m
12.8m
2.4m
satellite potential
φcg≈φs≈1~2V
Onset ofsubstorm
differential voltageΔ ≈V kV
coverglass potential
Potential
time
φcg
φs
GEO Plasma Environment Test Chamber
• Operational since March 2002• Diameter 60cm, Length 90cm• Minimum Pressure 3x10-7Torr• Equipped with an electron beam gun, a non-contacting surface potential probe, a Kaufman type plasma source, a xy-stage,
a low-pressure mercury lamp, halogen heat lamps, a video-image analysis system
Electron beam (max30kV)
Trek probe
X-Y stage controller
Trek probe
Solar array coupon
X-Y stage
LEO Plasma Environment Test Chamber
•Operational since 1998•Diameter 1m, Length 1.2m (excluding sub-chamber)•Minimum pressure 1x10-6Torr•Plasma density 1x1012m-3 or above•Equipped with an ECR plasma source, a deuterium UV lamp, a mass spectrometer, a x-y stage, a video-image analysis system, a high speed data acquisition system, halogen heat lamps, a spectrum analyzer, a metal-halaide lamp
Pressure chamber(1 MPa)
= 600 mm
time
time
Vfo
density
temperature
temperature
density
Vfo
Measurement of gas properties Measurement of flashover voltage
Build-up of database
Distribution of charge exchange ions and neutralizer electrons near a GEO satellite with an ion thruster
Ion density Electron density
Example of three-dimensional simulations
x
y
z rB
rvd
0 20 40 60
electron density
Magnetic field
plasma flow
