The document summarizes the design of a space-based solar power system. It proposes using five satellites in geosynchronous orbit to collect and transmit solar power to five ground stations via microwave beams. Each satellite would have a large solar concentrating structure and transmitter to beam 1 GW of power to Earth. The design aims to meet requirements to transmit 1 GW of power by 2040 for under $21 billion.

1. Solar POwer Collector
Final Design Review
1

2. Brittany Barker
Lead Systems Engineer
2

3. Overview
 RFP Requirements
 Basic Design
 Mission Timeline
3

4. RFP Requirements
Design a space based solar power system that can:
 Transmit 1 GW of power to Earth
 Be fully operational by 2040
 Cost under $21 billion
 Be safe and reliable
4

5. Basic Design
Transmitter
Solar Concentrating Structure
Connecting Trusses
Solar Array
5

6. Basic Design  Five satellites and five
ground stations
 Each satellite in GEO
 Mirrors always facing the
sun
 Parabolic shape
 Center cut out since solar
panels block it
 5 to 1 concentration factor
 Copper transmission
cables from solar array to
transmitter
 Rotating Transmitter
 2 Communications buses
6

7. Dimensions
 Solar Concentrating System
 Radius: 461 m
 Cutout Radius: 186 m
 Depth: 137 m
 Cutout depth: 21 m
 Solar Array
 Radius: 186 m
 Distance from mirrors: 388 m
 Transmitter
 Radius: 90 m
 Communications bus
 2.5 m x2.5 m x 1.5 m
7

8. Mass Estimates
Component Total Mass (kg)
Communications 120
SpaceBots 960
Copper Transmission Cable 13,500
Structure 31,300
Thrusters/Propellant 22,300
Reaction Wheels/Motors 12,500
Solar Panels 179,000
Transmitter 13,400
Propellant (Satellite) 91,250
Magnetrons 24,000
Mirrors 146,600
Total for one satellite 537,370
Total for SPOC 2,686,850
8

9. Zachary Roberts
Structural Engineer
9

10. Overview
 Structural Requirements
 Connecting Structure
 Solar Concentrating Structure
 Frame
 Hexagonal Mirror Segments
 Interconnects/Assembly
Satellite dimensions. 10

11. Structural Requirements
 Confine subsystems to proper positions
 Protect subsystems from external forces
 Two primary components:
 Solar Concentrating Structure (SCS)
 Reflect incoming light onto solar array
 Connecting Structure
 Connect solar array and SCS
SCS
BACKBONE
11
SPOC satellite.
SCS
Connecting
Structure
Solar
Array

12. Outer Rim (Green)
Inner Rim (Purple)
Backbone Segments (Red)
Connecting Trusses (Blue)
Solar Array Frame (Black)
SCS Frame/Connecting Structure
 Outer and Inner Rims
 Backbone
 Connecting Structure
ATK SRTM deployed [3.1].
12
Structural Components

13. SCS Hexagonal Mirror Segments
 Stowed Phase (mirror modules)
 Deployed Phase
MYLAR FILM
TRUSS SEGMENTS
CASING
Mirror module (top casing removed).
Deployed hexagonal mirror segment.
13
MIRROR MODULE

14. SCS Interconnects/Assembly
 Mirror Segment Interconnects
 Assembly Procedure
Mirror segment interconnect concept.
Mirror segment
interconnect locations.
Mirror segment
rings and
curvature.
14

15. Derek Chen
Payload Power Engineer
15

16. Payload Power System Overview
16

17. Power Generation
 GaInP2/GaAs/Ge Triple
Junction Cells
 Concentrated at 5x
 Proposed AM0 Efficiency:
45%
A GaInP2/GaAs/GeTriple Junction
Cell
1. http://www.spectrolab.com/DataSheets/cells/PV%20XTJ%20Cell%205-20-10.pdf 17

18. Power Generation
2. Table adopted from Kazmerski, Lawrence, “Best Research-Cell Efficiencies” ,Rev 4-2011, National
Renewable Energy Laboratory, March 2011
0
10
20
30
40
50
60
2000 2010 2020 2030 2040
Efficiencies(%)
Year
Projected 3-J Conversion Efficiencies
AM1.5 Efficiency
5x AM0 Effiency
5x AM0 Effiency @ 60°C degree
18

19. On-Board Power Distribution
Solar
PVs
35V DC-DC
Convertor
Rotating
Slip Ring
Magnetron
Subarray
35 V
Solar
PVs
35V DC-DC
Convertor
Solar
PVs
35V DC-DC
Convertor
35 V
Solar
PVs
35V DC-DC
Convertor
35 V
.
.
.
Magnetron
Subarray
Magnetron
Subarray
.
.
.
Magnetron
Subarray
Antenna
Subarray
Antenna
Subarray
Antenna
Subarray
Antenna
Subarray
∑ 7 kV7 kV
19

20. Rotating Slip Ring
 Parallel 7 kV connection
Rotating Transmitter Core
Outer ShieldingConductive Base
Rotating Brushes +
Inner Core
20

21. DC-to-Microwave Generation
RF Source
Power/module/kg
(kW/kg)
Efficiency
(%)
Advantages Disadvantages
Klystron 1.84 75
High Voltage,
Simpler control due
to lowest unit count
Heavy,
Moderately
Expensive
GaN Solid-
State Power
Amplifier
10 70 Compact
Expensive,
Thermal Issues,
Complex
Control
Magnetron 5 87.5
Inexpensive,
Compact,
Phase Controllable
N/A
21

22. Magnetron Subarrays
 Operate at 7 kV
 Efficiency: 85.5%
 Can output 5 kW
 Phase controllable to 1° RMS
 Lightweight and compact
22

23. Antenna Array
 Slotted Waveguide Antenna
 Antenna Gain: 13 dB
3. http://www.radartutorial.eu/06.antennas/pic/phased_array_(2).print.jpg
23

24. SPOC Efficiency Chain
Component Efficiency Power After Loss (GW)
Mirror Area Solar Flux @ 1367 W/m2 3.85
Mirrors 0.9300 3.582
Solar Panels AM0 5x @ 60°C 0.4760 1.705
Solar Panel EOL Efficiency 0.9460 1.613
Onboard Conductors & Wires 0.9500 1.532
35 MW for Spacecraft Power 35 MW 1.497
Slip-ring Rotating Brush 0.9995 1.496
Magnetron DC to Microwave 0.8750 1.309
Antenna 0.9700 1.270
Free Space Transmission 1.000 1.270
Atmospheric Transmission 0.9800 1.245
Ground Receiver 0.8800 1.095
Downtime and Backup Buffer 0.9132 1.000
Total 0.2658 1.000
24

25. SPOC Payload Power Summary
Item Specifications
Photovoltaics Total area: 523,780 m2
Number of Panels: 261,890
Efficiency: 45% @ AM0
Operation temperature: 60°C
Transmission
type
Microwave beam
Bandwidth: 150 MHz
Frequency: 5.8 GHz
Up-converter Total units: 185,000
Copper Wiring Total volume: 67,556.61 m3
Magnetrons Total units: 240,720
Antenna Material: Graphic composite
Aluminum
Size: 90 m (radius) per satellite
25

26. Steven Turner
26

27. Overview
 Refueling satellite placed in LEO
 Payloads launched into LEO
 Orbital Transfer Vehicle docks with and carries
payload to GEO on a spiral transfer
 Reaction wheels counteract rotating transmitter
 Hall Effect thrusters provide station-keeping
27

28. European Space Agency’s Automated Transfer Vehicle
Modification Specifications
Mass 13,336 kg
Payload Up to 46,000 kg to GEO
Attitude and Control Thrusters Busek BHT-600 Hall Effect Thrusters
Main Engines High Isp Nested Hall Effect Thrusters
Nuclear Reactor SAFE-400 Space Fission Reactor
Propellant Up to 5,495 kg of Krypton
28

29. Refueling Satellite
 Pentagonal prism hub with five propellant tanks docked to it using
ATV docking system
 ATVs dock with propellant tanks to refuel
 Each tank carries 96,000 kg for a total of 480,000 kg
 Located in 1,000 km circular orbit at zero inclination
 Tanks coated in optical solar reflectors to reduces pressure to minimize
leakage
29

30.  Spiral transfer
 Time of flight varies for mass of the payload
Payload Mass (kg) Round Trip Time (days)
Propellant 45,612 163
Magnetrons 24,002 115
Solar Panels/Transmitter/Reaction Wheels 25,550 119
Solar Panels 37,577 145
Mirrors 42,680 156
Transmitter 4,618 73
Structure/Robots/Communications/Thrusters 30,000 128
Power Cable/Mirrors 37,214 144
Refueling Station Propellant Tank 98,500 99 (no return trip)
•Total time of flight to transfer all payloads to GEO is 6.7 years
•Time of flight and propellant requirements verified with STK; maximum error within
0.8%
•Total propellant required is 478,470 kg of krypton 30

31. Orbit Transfer
31

32. Station-Keeping
 10 reaction wheels counteract angular momentum
from spinning reaction wheels
 45 NASA-300M Hall Effect thrusters perform station-
keeping and de-spin reaction wheels
 Delta-v of 128 m/s per year, lifetime 30 years
32

33. Thruster and Reaction Wheel Location
33

34. Station-Keeping
Maneuver Burn Time Delta-V per burn
De-spin reaction wheels 47.3 minutes every 52.1
days
0.0109 m/s
E-W station-keeping 56 minutes, twice per day
at 6AM and 6PM
0.14 m/s
E-W station-keeping 5.7 minutes at midnight
every 60 days
0.006 m/s
N-S station-keeping 16 hours at 6AM every 60
days
2.5 m/s
•Station-keeping propellant mass is 57,770 kg per satellite
•Thrusters magnetically shielded to greatly reduce erosion (approx. 600 times
less) [4.1]
•Each satellite will carry 60,000 kg of propellant for station-keeping during
construction
34

35. Aliya Burkit
CC&DH Engineer
35

36. Overview
 General design
 Command and Data Handling
 Communications
 Link budget
 Refueling satellite
 ATV
36

37. General design
Outer rim truss
Antenna
Star
trackers
Sun
sensors
Direction
to Sun Sun
sensor
s
Solar
panels
2.4 m dia
2.5 m
2.5 m
1.25 m
37

38. General design
38

39. Command and Data Handling
 RAD750 single-board computer by BAE Systems Electronic
Solutions
 Altair HB+ star tracker by Surrey Satellite Technologies
 SSOC-A60 sun sensor by Innovative Solutions In Space
 Thrusters and reaction wheels control
39

40. Communications
 Ku-band
 12.5 GHz downlink
 14.0 GHz uplink
 Ground station in Jupiter, Florida + backup
 Frequency division multiple access (FDMA)
40
Google maps

41. Communications
 Spacecraft antenna
 2.4 m diameter parabolic
reflector offset feed by Av
Comm
 deployable
• Ground station
antenna
– 2.4 m diameter parabolic
reflector offset feed by
Skyware Global
[5.2][5.1]
41

42. Link Budget
 Data rate = 12 kbps
 4kbps - commands
 8 kbps - health and telemetry data
 Margin = 34.72
 enough to account for rain attenuation
42

43. Refueling satellite
 Same components as for SPOC satellites
Sun sensors
Antenna
43

44. ATV communication
 Modified ATV as orbit transfer tug
 NASA TDRSS network for continuous communication
44
[5.3]

45. Braven Leung
Power/Thermal Systems Engineer
45

46. Overview
 Power Management and Distribution
 Satellite Power Consumption – Collection
 Satellite Power Consumption - Eclipse
 Thermal Trade Study
 Solar Concentration Factor Selection
 Photovoltaic Thermal Control System Configuration
 Thermal Control System Summary
46

47. Power Management and Distribution
Electrical Power System Schematic
• High Voltage,
Long Distance
47

48. Satellite Power Consumption - Collection
*** Not an actual power mode for satellite, validation purposes only
Component Quantity
Power Requirement For Mode (W)
Radiation
Pressure
Control
Burn
Longitudinal
& Inclination
Control Burn
Reaction
Wheel De-
Spin Control
Burn
Max Possible
Power
Draw***
Star Trackers 4 17 17 17 34
Sun Sensor 6 1 1 1 2
Spacecraft Antenna 2 38 38 38 76
Onboard Computer 2 5 5 5 10
CC&DH Total 61 61 61 122
Electric Thrusters 45 240,000 520,000 60,000 900,000
Reaction Wheels 10 3,500,000 3,500,000 3,500,000 3,500,000
Propulsion Total 3,740,000 4,020,000 3,560,000 4,400,000
Total Power 3,740,061 4,020,061 3,560,061 4,400,122
48

49. Satellite Power Consumption – Eclipse
Component Quantity Power (W)
Star Trackers 4 17
Sun Sensor 6 1
Spacecraft Antenna 2 38
Onboard Computer 2 5
CC&DH Total 61
Electric Thrusters 45 -
Reaction Wheels 10 -
Propulsion Total 0
Total Power 61
49

50. Thermal Trade Study
Solution
Technology
System
Type
Description Issues Conclusion
Pure
Conduction
PTCS
Conductive material to dissipate
heat
● Heat flux too high
● Likely to melt
Impractical
Heat Pipes PTCS
Uses pipes with fluids that cool PV
cells with vapor/condensing
● Small scale
● Only low thermal loads
Infeasible
Fluid Loop ATCS
Pumps system with fluid to
remove heat
● Expensive to launch
● Potential leaks
Impractical
Passive
Radiation
PTCS
Passively managed with paints and
coatings
● Low conc. factor
Viable, for low
conc. factor
50

51. Solar Concentration Factor Selection
Solar concentration factor over
temperature range to determine
optimal Photovoltaic Thermal
Control System configurations.
• Static Heat Balance
• Concentration Factor of 5 : 1
• PV Cells Temp ≈ 60 ̊C
-250 -200 -150 -100 -50 0 50 100 150 200 250
0
5
10
15
20
25
30
Solar Concentration Factor Over Temperature Range
Solar Collector Temperature (ºC)
SolarConcentrationFactor
X: 62
Y: 5.003
▪ Design Point
–– Optimal PV Regime
(Currently)
51

52. Photovoltaic Thermal Control System
Configuration
52

53. Thermal Control System Summary
Location Component
Emissivity
(ɛ)
Function
Solar Panel
Rear (Sun-
Facing)
Optical Solar
Reflectors
0.92
High emissivity surface component
Reflect incoming solar rays
Facilitate thermal radiation
Solar Panel
Front Main
(Reflector-
Facing)
PV Laminate 0.75
Emissivity surface coating
Facilitate thermal radiation
Reflector Rear
Face
None 0.53
The Mylar film reflector will
naturally radiate thermal loads
without any PTCS assistance
53

54. Thermal Control System Summary
Optical Solar Reflector PV Laminate None
PV - Sun Facing Reflector BackPV - Reflector Facing
* Locations highlighted in yellow
54

55. Emily Zavala
Ground Power Engineer
55

56. Overview
 Introduction
 Rectenna Elements
 Ground Station Design
 Ground Station Locations
56

57. Introduction
 Five ground power receiving stations (GPRS)
 Rectenna design that is safe, cost effective, and least
obstructive to the general population
57

58. Rectenna Elements
 Serrated Panel
 Ground Plane Mesh
 Dipole Foreplane
Depiction of diode foreplane [7.1]
58

59. Ground Station Design
Diagram of a panel indicating sub-panel and strand
59

60. Ground Station Design
Support structure composed of bar joist frame and I-beams with concrete footing.
60

61. Ground Station Design Cont.
Visualization of a ground power receiving station [7.2]
61

62. Ground Station Locations
Federal Lands Nature Topography Weather
National Recreation Areas Land and water Open mountains Windstorms
American Indian
Reservations
National Forests Hills Hail
Military Reservations Marsh Vegetation Mountains Thunderstorms
Other Federal Lands Wetlands Sheet Rainfall
Endangered species habitat Acid Rainfall
62

63. Map of unacceptable areas for rectennas [7.3]
63

64. Map of unacceptable areas for rectennas including 40º latitude [7.3]
64

65. Map of unacceptable areas for rectennas with seismic hazard consideration
65

66. Map of acceptable GPRS locations
66

67. Selected GPRS Locations and existing power plants and power lines [7.3]
67

68. Lillian Helms
Launch Systems Engineer
68

69. Overview
 Launch Vehicle Selection
 Payload Packing Logistics
 ATV
 Launch Timeline
69

70. Launch Vehicle Selection
Launch Vehicle Company Launch Site(s) Payload to LEO
(kg)
Cost per Launch ($
in millions)
Cost per kilogram
to LEO ($/kg)
Atlas V HLV ULA CCAFS, VAFB 29,400 130 4,400
Delta IV Heavy ULA CCAFS, VAFB 22,560 140 6,200
Falcon Heavy SpaceX CCAFS 53,000 128 2,400
Falcon 9 SpaceX CCAFS 13,150 54 4,100
Ariane 5 ArianeSpace CSG 21,000 120 5,700
Falcon Super Heavy SpaceX CCAFS 150,000 300 2,000
70

71. Payload Fairing Packing
Top Views of Mirrors/
Solar Panels/Transmitter
Payload fairings 1-3.
Thrusters/R
obots
(X2)
71

72. Payload Fairing Packing
Reaction Wheels
Payload fairings 4-6.
72

73. Payload Fairing Packing
2.9 m
Payload fairings 7-10.
73

74. Tug (ATV)
Payload Number Number of Tugs Needed Components Total Mass (kg) Time of Flight (round
trip, in days)
1 1 Propellant (ATVs) 98,500 99
2 2 Robonauts/
Communications/Structure/
Thrusters/Propellant
60,000 128
3 2 Solar Panels 75,150 145
4 2 Solar Panels 75,150 145
5 2 Solar Panels/
Transmitter/
Reaction Wheels
51,000 119
6 1 Transmitter 9,300 73
7 1 Magnetrons 24,000 115
8 2 Propellant (Satellite) 91,250 163
9 2 Power Cable/Mirrors 74,500 144
10 2 Mirrors 85,400 156
74

75. ATV Docking
 Used on current ATV
 Equipped with guidance
system to direct ADA
into PDA
 ATV has ADA
 Refueling satellite has 5
PDAs
 Payload containers have
ADA and PDA (one on
each end)
3rd IAASS Conference - 4 -
“Building a safer space together”
ATV to ISS docking using th
ISS docking
port: PDA
ATV Docking
system: ADA
cone
socket
Grooves &
Latch stops
Probe
head
latches
Docking system [8.9]
75

76. Launch Timeline
Launch and assembly progression
76

77. Launch Timeline
 56 total launches (6 Ariane 5, 50 Falcon Super Heavy)
 First satellite with ATVs total of 426 days (2031) to
launch
 Satellites 2-5 average of 392 days to launch per satellite
 Total of 6.4 years to fully launch all satellite
components, 6.7 (late 2036) years for all components
to be in GEO
77

78. Brittany Barker
Lead Systems Engineer
78

79. Risk Assessment
Number 1 2 3 4 5
Probability of
Occurrence
Very low Low Moderate High Very High
Not expected
to occur
Less likely
than not to
occur
May or may
not occur
More than
likely to occur
Expected to
occur
Impact on
Mission
Negligible Minor Moderate Significant Severe
Little to no
impact
Slight impact
but system still
functional
close to
intended
Medium
impact;
System
functional as a
whole but less
than intended
Parts of
system not
functional;
Loss of
mission
objective
Total system
and mission
failure
Total Risk Weight = Probability of Occurrence x Impact on Mission
79

80. Overall Risk
Probability
5 Very High 1
4 High
3 Moderate 2 1
2 Low 2 7 6 3 2
1 Very Low 4 1 2
Very Low Low Moderate High Very High
1 2 3 4 5
Impact
Mitigated risk by:
• Multiplicity
• Redundancy
• Buffers
• Factoring in additional costs
80

81. Cost Analysis
 Determined raw materials costs for each subsystem
 Applied Advanced Missions Cost Model (AMCM) for
items that needed to be research and developed
further
 Applied cost estimating relations (CER) for parametric
cost estimation for items that needed to be
manufactured after raw materials were purchased
81

82. Project SPOC Costs
Subsystem Total Cost ($FY13)
Structures 29,284,800
Power Generation and Transmission 812,770,000
Orbit Transfer, Attitude
Determination, and Station Keeping 1,769,129,000
CC&DH 12,575,000
Spacecraft Power and Thermal
Systems 22,615,800
Ground Power Receiving Station 2,079,521,900
Launch Services 13,440,000,000
Research and developmentcosts 2,198,250,000
Manufacturing costs 18,937,270
Total 20,383,000,000
82

83. Questions?
83

84. Structure Appendix
84

85. 85
Truss
Axial Strength
(N)
Column Length
(m)
Mass per
Length (kg/m)
Truss Radius
(m)
Deployment
Method
ILC Dover
UltraBoom
1413 9.06 0.1445 0.09 Inflatable
L’Garde SSP
truss
2473 78.28 0.7 0.68 Thermoset
ATK GR1 494 40.30 0.07 0.197 Uncoiled
ATK SRTM 5270 60 3.835 0.561 Articulating
Deployable Truss Comparison [3.2]
Thin-Film Comparison [3.3]
Thin Film
Density
(kg/m3)
Tensile
Strength
(MPa)
Thickness
(µm)
Kapton 1,420 138 12.7
CP1 1,434 99.97 25.4
Mylar 1,390 179.3 12.2

86. 86

87. Component Dimensions
Mirror module (377/satellite)
Mirror Truss
1.1 m side length (hexagon)
388 kg
3.14 m2
25 m length (0.75 m stowed)
0.25 m diameter
Thin film reflector segment 25.35 m side length (hexagon)
1640 m2
Frame truss (136/satellite) 60 m length
1.121 m diameter
3.85 kg/m
Backbone Segments (4) 300 m each
Connecting Trusses (4) 375 m each
Overall Structure
Outer Rim
Inner Rim
Radius – 461.2 m
Depth – 137 m
Cutout Depth – 21.5 m
Solar Array Radius – 182.6 m
Focal Length 388.15 m
2900 m circumference
1150 m circumference 87

88. Payload Power Appendix
88

89. Antenna Power Density
89

90. Best Research-Cell Efficiencies
2. Table from Kazmerski, Lawrence, “Best Research-Cell Efficiencies” ,Rev 4-2011, National Renewable
Energy Laboratory, March 2011 90

91. Magnetron Subarray Density
Step % of Max
Power Density
Subarrays Magnetrons/
Subarray
Number of
Magnetrons
Outer Radius
of Step Ring
(m)
1 100% 19 443 8610 24.879
2 83.33% 24 369 8782 37.101
3 66.67% 19 295 5741 44.671
4 55.56% 24 246 5854 52.469
5 44.44% 31 197 6038 61.071
6 33.33% 31 148 4528 28.602
7 25% 12 111 1391 71.457
8 22.22% 31 98 3019 77.992
9 16.67% 43 74 3191 86.367
10 11.11% 20 49 990 90
Total 255 48144
91

92. Solar Cell Band Gaps
Substrate Band Gap (eV) Max.
Wavelength
(nm)
GaInP2 1.8 689.28
GaAs 1.43 867.62
Ge 0.67 1851.79
92

93. Solar Spectrum
93

94. ATV Appendix
High Isp NHT engine (per unit) [4.2] Isp: 5000s
Thrust: 6.1 N
Efficiency: 65%
Power consumption: 200 kW
Mass: 280 kg
Busek BHT-600 thruster (per unit) [4.3] Isp: 1585 s
Thrust: 42 mN
Efficiency: 49%
Power consumption: 600 W
SAFE-400 Fission Reactor (per unit) [4.4] 400 kWt
200 kWe
540 kg
Krypton Propellant ATV Maximum 5,495 kg
94

95. ATV Appendix
Payload Propellant Round Trip (kg)
Propellant 5,493
Magnetrons 3,526
Solar Panels/Transmitter/Reaction Wheels 3,663
Solar Panels 4,763
Mirrors 5,228
Transmitter 1,761
Structure/Robots/Communications/Thrust
ers
4,073
Power Cable/Mirrors 4,730
Refueling Station Propellant Tank 4,091 (no return trip)
95

96. Payload Round Trip Time (days)
Propellant 163
Magnetrons 115
Solar Panels/Transmitter/Reaction Wheels 119
Solar Panels 145
Mirrors 156
Transmitter 73
Structure/Robots/Communications/Thrust
ers
128
Power Cable/Mirrors 144
Refueling Station Propellant Tank 99 (no return trip)
ATV Appendix
96

97. Refueling Satellite Appendix
 5 Titanium 6Al-4V fuel tanks 2 mm thick, interior volume
530 m3 carry 96,000 kg krypton each
 Spacecraft bus is pentagonal prism with a PDA on all 5
sides for propellant tanks to dock with
 Tanks will be coated with optical solar reflectors to keep
maximum temperature at 150 K
 Orientation will remain in the plane of the ecliptic to
minimize surface area exposed to the Sun
Busek BHT-1000 [4.3] Isp: 1,750s
Thrust: 58 N
Power consumption: 1 kW
Mass: ~5 kg
97

98. Station-Keeping Appendix
NASA-300M thrusters [4.5] Isp: 3,220s
Thrust: 1.13 N
Efficiency: 63%
Power consumption: 20 kW
Mass: ~15 kg
Reaction Wheels Aluminum
0.75 m radius
0.23 m height
3,500 RPM
98

99. References
 [3.1] “Articulated Mast Systems,” ATK, [http://www.atk.com/wp-
content/uploads/2012/09/ADAM-2011.pdf. Accessed 2/12/13.]
99

100. References
 [4.1] Mikellides, Ioannis G., Katz, Ira, Hofer, Richard R., and Goebel,
Dan M., “Magnetic shielding of walls from the unmagnetized ion beam
in a Hall thruster”, Applied Physics Letters, Jan 2013.
 [4.2] Space Mission Analysis and Design, Microcosm Press, CA, pp 560.
 [4.3] “Low Power Hall Effect Thrusters”, Busek, 2012.
[http://www.busek.com/index_htm_files/70008510_revA.pdf. Accessed
02/15/2013].
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103