3. Outline and Contents
1. Introduction of Power Systems
2. Energy Sources
3. Energy Storage Devices
4. Control of Power System and Distribution
[Online Quiz] by 18/Oct (Monday)
4. Learning Outcomes on Ch1
• Able to answer the following questions
⎻ What are the roles of Power System?
⎻ What is the design principle of Power System?
⎻ How does the orbit of a spacecraft affect on the Power System design?
⎻ What are the roles of Primary Power source?
⎻ What kind of primary power source are available?
• Able to determine the dimension of a solar array
• Able to account the temperature and radiation effects on
the design of a solar array
• Able to determine the required capacity of batteries
6. 1. Introduction of Power System
• Generate, store and distribute electrical power
⎻ Supply a continuous source of electric power to spacecraft
loads during mission life.
⎻ Support power requirements for average and peak electrical
power.
Function
Power System Module (SSTL) Global Precipitation Measurement (NASA Goddard)
7. 1. Introduction of Power System
• Regulate voltage, protect against shorts, provide
electrical grounding
⎻ Provide converters for AC and regulated DC power buses.
⎻ Control and distribute electrical power to spacecraft
⎻ Provide command and telemetry capability for EPS (Electric
Power System) health and status as well as control by ground
station and/or an autonomous system.
⎻ Protect the spacecraft payload against failures within EPS
Function
8. 1. Introduction of Power System
1. Primary Power Source (Power Source):
2. Secondary Power Source (Energy Storage):
Key Components
• Converts “fuel” (can be solar radiation) into electrical power.
Examples:
Solar arrays, fuel cells, Radioisotope Thermal Generators (RTGs),
fission reactors and batteries
• Required to store energy and deliver electrical power to the
spacecraft when the primary power source is unavailable (e.g. solar
arrays during an eclipse).
Examples:
• Batteries and fuel cells can serve as this element.
9. 1. Introduction of Power System
3. Power control and distribution network:
Key Components
• Regulate current and/or voltage as the primary and secondary
power systems characteristics change
11. 1. Introduction of Power System
• Most important sizing requirements are
⎻ the demands for average and peak electrical power
⎻ the orbital profile (inclination and altitude)
• Must identify the electrical power loads for mission
operation at Beginning Of Life (BOL) and End Of
Life (EOL)
Sizing Requirements
12. 1. Introduction of Power System
• Average electrical power
• Peak electrical power
Effects of Mission Requirements
- The power generation system
Example: number of solar cells, primary size batteries
- The energy storage system given the eclipse period and depth of
discharge
- The energy storage system
Example: number of batteries, capacitor bank size
- The power-processing and distribution equipment
13. 1. Introduction of Power System
• Mission life
• Orbital parameters
Effects of Mission Requirements
- Longer mission life (> 7 years) implies
§ Extra redundancy design
§ Independent battery charging
§ Larger capability batteries
§ Larger arrays
- Defines incident solar energy, eclipse/Sun period, and radiation
environment
14. 1. Introduction of Power System
• Spacecraft configuration
Effects of Mission Requirements
- Spinner typically implies body-mounted solar cells
- 3-axis stabilised typically implies body-fixed and deployable solar
panels
16. 2. Power Source
• Power source generates primary electrical power
within the spacecraft.
• Primary source of power generate power over many
cycle (to support loads and charge batteries)
Examples
⎻ Solar (photovoltaic or solar cells)
⎻ Nuclear/Static power sources
⎻ Chemical/Dynamic power sources
⎻ Fuel cells (manned space missions)
Role of Power Source
17. 2. Power Source
1. Solar (photovoltaic or solar cells)
⎻ Most common for Earth-orbiting spacecraft
⎻ Convert solar radiation directly to electrical energy
Primary Power Sources
Solar array of Juno spacecraft (NASA) Solar array of Cloryspacecraft (NASA)
19. 2. Power Source
2. Nuclear/Static power sources
⎻ Heat source: plutonium-238 or uranium-235 (nuclear
reactor) for direct thermal-to-electrical conversion
Primary Power Sources
RTG of Mars Science Lab (Idaho National lab) RTG of New Horizons (NASA)
21. 2. Power Source
3. Chemical/Dynamic power sources
⎻ Heat source: concentrated solar radiation, plutonium-238,
or enriched uranium to produce electrical power using
Brayton, Stirling or Rankine cycles.
Primary Power Sources
Stirling generator (NASA)
26. 2. Power Source
Power Source Selection
• For short missions (e.g. launch)
⎻ batteries are appropriate
• Medium duration mission (e.g. manned)
⎻ commonly use fuel cells
• Low power, long duration missions
⎻ use solar arrays most often
• High power, long duration mission
⎻ Nuclear devices (RTGs)
27. • Convert solar radiant energy to electricity via the photovoltaic effect
• Power production depends on the intensity of solar radiation
A. Solar Array
2. Power Source
Solar Radiation
• Intensity Js of solar radiation at distance D from the Sun
2
2
W/m
4
S
P
J
D
p
é ù
= ë û
where P=3.8x1026 W, the total power
output of the Sun
(1.1)
Example:
Solar intensity at 1 A.U. ≈ 1350 W/m!
28. • The power of radiation falling on a surface is the product of JS
and the projected normal area, As, of the surface in the
direction of the Sun:
A. Solar Array – Solar radiation
2. Power Source
Sun
A AS
radiation
𝜽
Solar energy available at the surface, Isc:
I!" = 𝐽# ⋅ 𝐴# =
𝑃
4𝜋𝐷$
⋅ 𝐴 ⋅ 𝑐𝑜𝑠𝜃 (1.2)
29. Basic Configuration
• Each solar cell has a semiconductor p-n junction
A. Solar Array
2. Power Source
GaInP
GaAs
Ge
n
p
n
n
p
p
Tunnel
junctions
Metal
Front metal
GaInP
GaAs
Ge
n
p
n
n
p
p
Tunnel
junctions
Metal
Front metal
⁃ p-type material (silicon doped with boron): electron deficient
⁃ n-type material (silicon doped with phosphorus): electron excess
30. Operating Principle
• Radiation is converted to a potential across the cell with usable
electrical power
A. Solar Array
2. Power Source
Isc
Voc
[A] I_SA
V_SA
[V]
Voc
PHOTOVOLTAIC CELL MAX VOLTAGE
PHOTOVOLTAIC
CELL
MAX
CURRENT
IMPP
VMPP
[W] P_SA
V_SA
[V]
IMPP
VMPP
0W
0W
MPP
(maximum Power to load
R_LOAD (P_SA)
”SA”
SOLAR
ARRAY
I_SA
+
-
V_SA
31. • Si solar cells are rectangular, approx. 4cm by 7 cm, have a
conversion efficiency of 12~14% and have a thickness of
between 50 and 250 𝜇𝑚
A. Solar Array – Operating Principle
2. Power Source
32. • The solar array is neither an ideal voltage nor an ideal
current source.
• The nature of a solar array depends on the operating point P
and the excursion that we assume to impress to the point P in
the function of the load change and specific power
conditioning technology considered
A. Solar Array – Operating Principle
2. Power Source
33. • Radiation damage is a major problem for solar arrays
⎻ A glass cover is used to provide environmental and radiation
protection
⎻ To form arrays, cells are mounted on a substrate that may be
flexible (Fiber-glass reinforced Kapton) or rigid (carbon fiber
reinforced Kapton)
A. Solar Array
2. Power Source
34. A. Solar Array
2. Power Source
Array Type Power (kW) W/kg W/m2
XMM Rigid 2.5 32 215
Astra 2B Rigid 9.2 52 409
Comets Flexible 6.3 34 146
ISS(Wing) Flexible 32 29 135
Solar Array Performance Figures in Earth Orbit
36. Required Solar Array Power
• Need include the power to charge battery
A. Solar Array
2. Power Source
𝑃%& = 𝑃' 1 +
𝑉()
𝑡()
⋅
𝑡*+#
𝑉*+#
⋅ 𝑅𝐹 (1.3)
Psa = solar array power required, W
Pl = the load power
Tch = battery charge time, min
Tdis = battery discharge time, min
Vch = avg. battery charge voltage, V
Vdis = avg. battery discharge voltage, V
RF = battery ampere hour recharge fraction ratio
ηppt = peak power tracker converter efficiency
ηtr = peak power tracking accuracy
ηdr = discharge regulator circuit accuracy
ηcr = charger circuit accuracy
where
37. Solar cell (array) efficiency
i. Temperature
A. Solar Array
2. Power Source
Example:
40. iii. Sun-angle
• Total Efficiency,
A. Solar Array – Solar cell efficiency
2. Power Source
𝜂#,- = 𝑐𝑜𝑠𝜃
Sun
A AS
radiation
𝜽
𝜂./.0'
𝜂./.0' = 𝜂.123 𝜂40* 𝜂#,-
41. Determine the required size of a Solar Array
Recall: Solar Array Sizing in SESA2024 Astronautics
A. Solar Array
2. Power Source
𝐴%& =
𝑃567
𝑆 ⋅ 𝑐𝑜𝑠𝜃 ⋅ 𝜂 ⋅ 𝜂3 ⋅ (1 − 𝐷)
42. • Model a solar array
⎻ Each solar cell can be modelled as a battery
A. Solar Array – Sizing
2. Power Source
………..
……….. n
m
43. i. Determine number of cells, n
A. Solar Array – Sizing
2. Power Source
𝑛 =
𝑉&
𝑉
0
(1.4)
𝑉& = Required min. voltage for a solar array
𝑉
0 = Voltage for a single solar cell
Required Solar Array Voltage, VA
- 𝑉& ≥ 𝑉(),9&: + 𝑉*4/3
44. A. Solar Array – Sizing
2. Power Source
Voltage of a single solar cell, Va
- Need to consider: degradation and temperature effects
Use temperature coefficient at BOL
Use degradation factor
𝑉
0 = 𝑉
23 ⋅ 𝑏 + 𝑎 ⋅ (𝑇20; − 𝑇.04<1.) (1.5)
45. ii. Determine number of strings, m
⎻
⎻
⎻
A. Solar Array – Sizing
2. Power Source
𝐼#.4+-< = 𝐴 ⋅ 𝐽#.4+-<
𝑃#.4+-< = 𝐼#.4+-< ⋅ 𝑉#.4+-< W
𝑚 =
𝑃%&
𝑃#.4+-< − 𝑃33.
(1.6)
46. ii. Determine number of strings, m
⎻ Use a cell current (Ja [mA/cm2])
⎻ Need to consider: degradation + temperature + distance to sun
A. Solar Array – Sizing
2. Power Source
Use temperature coefficient at BOL
Use degradation factor
𝐽0 = 𝐽23 ⋅ 𝑏= + 𝑎= ⋅ 𝑇20; − 𝑇.04<1. (1.7)
47. Q. What do we need to consider on the design of
solar cell array?
A. Solar Array – Sizing
2. Power Source
• Orbital parameters
• Bus voltage
• Power requirement (payload / bus / battery)
• Operating conditions (temperature / radiation)
• Magnetic field (magnetometers)
48. • Batteries are the only secondary power supply that have been used
• They provided power when the primary source is not available i.e.
during eclipse when used as a back-up for solar arrays
• Batteries must be recharged when the primary source is available,
in Earth orbit this results 5000 - 6000 charge/discharge cycles per
year
Introduction
3. Energy Storage
49. • Individuals cells connected in series to increase the voltage or in
parallel to increase the current.
• The energy stored in the batteries is the watt-hour (Wh) or ampere-
hour (Ah) capacity.
• Cell cycle life, specific weight (kWh/kg) and volume (kWh/m3)
affect acceptability of particular batteries
• Historically Ni-Cd and Ag-Zn batteries are used in LEO and Ni-H2
batteries are used in GEO
• Battery capacity degrades with time, but some can be regained by
reconditioning (full discharge before charging)
Introduction
3. Energy Storage
50. • Considerations
⎻ Size, weight, configuration
⎻ Duty cycle / Number of cycles
⎻ Limits on depth-of-discharge
⎻ Activation time and storage time
⎻ Voltage and current loading
⎻ Cost, reliability, maintainability and
producibility
Introduction
3. Energy Storage
Type Specific energy (Wh/kg)
Ni-Cd 39
Ni-H2 52
Ag-Zn 60
Ni-MH 60
Li-Ion 80
Li-TiS2 125
Na-S 150
Battery Performance
51. Example:
Hubble space telescope (HST) and Intelsat VII Ni-H2 battery data
summary
Introduction
3. Energy Storage
Parameter HST Intelsat VII
Specific energy (W.h/kg) 57.14 61.26
Capacity (Amp.h) 96 91.5
Cell dimensions:
diameter (cm) 9.03 8.89
length (cm) 23.62 23.67
terminal-terminal (cm) 24.66 29.67
Cell Mass (kg) 2.1 1.867
52. 3. Energy Storage
• Primary battery cells
⎻ Convert chemical energy into electrical.
⎻ Cannot reverse conversion (no recharge)
• Secondary battery cells
⎻ Convert chemical energy into electrical during discharge.
⎻ Convert electrical energy into chemical during charge.
Introduction
53. 3. Energy Storage
NOTE:
• Li-Ion Battery charge/discharge management:
⎻ Limiting discharge depending on mission calendar and cycling
requirement
• Low number of charge/discharge à Higher DoD allowed (up to 70% in
GEO)
• High number of charge/discharge cycles à Lower DoD allowed (30% in
LEO)
⎻ Precautions: Li-Ion technology do not tolerate
• Over-charge (typ 1-2% over the nominal end of charge voltage of 4.1 V-
4.2 V per cell)
• Over-discharge (typ 1-2% under the nominal end of discharge voltage of
2.5 V-2.9 V per cell)
Introduction
54. 3. Energy Storage
Most important figures of merit are:
⎻ W∙h efficiency: ratio of energy provided in discharge and the energy
accumulated over charge
⎻ Mass specific energy (W∙h/kg)
Charge and Discharge Cycles
t = 0 t = 36min t = 96min
100%
0
60%
80%
40%
20%
EXAMPLE: BATTERY CYCLING OF LEO MISSION
TIME
STATE
OF CHARGE
DoD ≈ 20%
t=0 t=1h t=24h
100%
0
60%
80%
40%
20%
EXAMPLE: BATTERY CYCLING OF GEO MISSION
TIME
STATE
OF CHARGE
DoD ≈ 70%
55. 3. Energy Storage
NiCd cells
• Using nylon as a separator.
• Used for energy storage on the
majority of spacecraft during the
first three decades of the space
era.
• Reasonable high energy density.
• Requires relatively simple charge
control systems.
58. 3. Energy Storage
NiH2 cells
• Can discharge to greater depth
than NiCd batteries.
• Essentially a pressure vessel.
• Sealed to withstand high internal
pressure during overcharge
(3.4 ~ 6.2 MPa)
• Electrolyte: an aqueous solution
of potassium hydroxide (31%
concentration by weight)
60. 3. Energy Storage
Li-ion cells
• Higher energy per unit weight
• Higher discharge voltage than Ni-
based chemistries (avg. 3.5 V)
• Very good storage life
è good for deep-space probes and
planetary landers (requires a limited
number of recharges and long mission
life)
• Simpler thermal control (Higher
efficiency: > 90%)
• No memory effect
• Cell voltage is a very good indicator
of the state of charge and is
minimally dependent on cell
temperature.
61. 3. Energy Storage
Li-ion cells
• No tolerance for overheating!!
• Poor performance at low temperature
• Cells cannot be assembled in the discharged state
62. 3. Energy Storage
• Number of cells in a battery:
Battery Capacity
𝑁 =
𝑉
0><
𝑉(1'' 0><
where:
Vavg = average battery bus voltage (system requirement)
Vcell avg = average battery cell voltage during discharge
N = the number of cells in series
63. 3. Energy Storage
• Computing battery capacity as a function of the required
power in eclipse and its characteristics:
Battery Capacity
𝐶4 =
𝑃1𝑡1
𝐷𝑜𝐷 ⋅ 𝜂
W ⋅ h (1.8)
Note:
• For finding a charge current, need to use Cr in Ah which can be
obtained by dividing with Charge voltage of a battery (NOT a
battery cell)
where:
Pe = the average eclipse load in Watt
te = the correspondent maximum eclipse time in hours
DoD = the limit on battery’s Depth-Of-Discharge
ƞ = transmission efficiency between batteries and load (around 0.9)
64. 3. Energy Storage
Example:
⎻ Pe = 110 W
⎻ te = 35.3 min/60=0.5917 h
⎻ DOD = 20% (upper limit)
⎻ n = 0.9
Battery Capacity
𝐶4 =
𝑃1𝑡1
𝐷𝑜𝐷 ⋅ 𝜂
W ⋅ h ∴ 𝐶4= 361.59 W ⋅ h
65. 3. Energy Storage
Requirements of Batteries for Spacecraft
The battery must operate reliably and unattended for a long
period in the space environment:
⎻ peak power delivery (NiH2 and NiCd particularly good)
⎻ rapid recharge at high rates (caution needed against overcharge)
⎻ large number of charge/discharge cycles (depends on depth of
discharge and battery chemistry)
⎻ high recharge efficiency (k-factor) (NiH2, NiCd ~80%; Li-ion ~95%)
⎻ good hermetic seal for operation in a vacuum
⎻ ability to survive launch (vibration, shock)
⎻ stable electrical characteristics
⎻ availability of electrically well-matched cells
⎻ ability to withstand over-charge (NiH2 and NiCd can, Li-ion cannot!)
⎻ Maximum specific energy or energy density (Li-ion is by far the
best)
⎻ high reliability
67. 4. Control of Power System and Distribution
• The power distribution consists of:
⎻ Harness (15-25% of the total EPS mass)
⎻ Fault protection and switch gear to turn on and off spacecraft loads
⎻ Command decoders to command specific loads
Introduction
Lunar Recognisance Orbiter harness Magnetospheric Multi-scale Spacecraft (MMS) spacecraft
harness
68. 4. Control of Power System and Distribution
• Power distribution architecture depends on the dimension
and complexity of the spacecraft and on the demands in
terms of power.
• Power distribution architecture:
⎻ Distributed: each load has its own dedicated feeding and control
system
⎻ Centralised: everything is controlled from the central bus
Introduction
69. 4. Control of Power System and Distribution
• Power regulation and control system is in charge of regulating:
⎻ Power distributed to the loads
⎻ Current and voltage of the bus
⎻ Charge/discharge cycle of the battery
• The bus voltage control can be:
⎻ Unregulated: load bus voltage varies significantly. Power bus
regulation derives from battery regulation
⎻ Quasi-regulated: subsystems regulate the bus voltage during
battery charge but not during battery discharge
⎻ Fully regulated: employs charge and discharge regulators
• Power and energy transfer to loads and battery can be performed by:
⎻ Direct Energy Transfer systems
⎻ Peak Power Tracking systems
70. 4. Control of Power System and Distribution
• The power from the solar array fluctuates dependent upon
illumination angle and eclipse
• The battery voltage fluctuates dependent upon state of charge
• The spacecraft sub-systems and payloads generally require
stable voltages and to be isolated from the effects of other
systems
• It is necessary to monitor and switch ON/OFF the spacecraft
loads
Why do we need power control and distribution system?
71. 4. Control of Power System and Distribution
Power Bus Concepts
REGULATED POWER BUS STANDARDS VOLTAGE
LEVELS
AND TYPICAL POWER LEVELS
UN-REGULATED POWER BUS STANDARDS
VOLTAGE LEVELS
AND TYPICAL POWER LEVELS
+ 28V ± 1%
Power up to ca 2 kW
(Usually: ESA SCIENCE spacecraft)
+17.5 .. +29.4V
Power up to ca 500 W
(Small and experimental Satellites)
+ 50V ± 1%
Power up to ca 8-10 kW
(Usually: TELECOM spacecrafts)
+20.0 .. +33.6V
Power up to ca 3 kW
(Usually: ESA EOP spacecrafts)
+ 100V ± 1%
Power > 5 kW
(Usually: TELECOM spacecrafts)
+22.5 .. +37.8V (or 36.9V)
Power up to ca 3 kW
(Usually: ESA EOP spacecrafts)
+ 120V ± 1%
Power ca > 8 kW
(This is only known from ISS)
+25.0 .. +42.0V
Power up to ca 5 kW
(Very unusual)
