YP in space2018_bcn_all_slides_theory_x2

1
IEEE YPinSpace 2018
© UPC‐BarcelonaTech ‐ 2018 1
IEEE Young Professionals in Space
… about this bootcamp
IEEE YPinSpace 2018
Organization
Morning: Invited Lectures (see web site)
Theory and Practical Sessions Alternate:
Theory Hands on session
1T Orbits 1P Orbits Lab using the Princeton Satellite Toolbox and
Orbitron
2T Space Environment and Thermal Control 2P Thermal Control Analysis using the Princeton
Satellite Toolbox
3T On Board Data Handling System (OBDH) 3P OBDH: Getting stated with Arduino
4T Attitude Determination and Control System (ADCS) 4P ADCS: Understanding Sun sensors, reaction wheels,
magnetorquers, GPS…
5T Electrical Power Supply System (EPS) 5P EPS: Understanding solar cells, MPPT, battery
chargers, batteries…
6T Telemetry, Tracking and Control System (TTC) 6P1 TTC: E2E communications and transceivers in radio
amateur bands
6P2 TTC: Ground Station: Manufacturing a Yagi antenna
and using the RTL to monitor the spectrum and decode
APRS packets
7P E2E tests, rocket launch, and telemetry tracking

2
IEEE YPinSpace 2018
Housekeeping information
Hands on sessions:
Electronics Eng. Labs
C4 building, floor ‐1
Theory Sessions:
Agora room
Plaça Telecos
Lunch
Unity Cafeteria
Plaça Telecos
Coffee breaks
B3 building – basement
Plaça Telecos
Lunch boxes will be
provided for Saturday
trip to Alcolea de Cinca
Unity Cafeteria
Plaça Telecos
Invited lectures
Registration 16/07/2018
Water rockets
side event
IEEE YPinSpace 2018
Satellite subsystems:
• Electrical Power Supply (EPS):
solar panels, battery charger,
and batteries
• On Board Computer (OBC):
CPU, memory
• Communications Systems (COMS)
transmitter(s) + antennas
• Attitude Determination and Control System (ADCS):
sensors + actuators (reaction wheels & magnetorquers)
• Payloads (P/L)

3
IEEE YPinSpace 2018
EPS
ADCS
OBDH
COMMS
P/L
SATELLITE
GROUND
SEGMENT
RTL‐SDR
(spectrum monitoring)
IEEE YPinSpace 2018

1
T1. Introduction and Orbits
IEEE Young Professionals in Space, Barcelona July 17‐21 2018 ‐ © UPC‐BarcelonaTech ‐ 2018 1
• Objectives of the lesson
- Learn the phases and segments of a satellite mission
- Learn about Orbits:
• Understand Kepler laws
• Understand Orbital Parameters
• Understand Perturbation factors
• Learn the different types of orbits and their applications
• Learn how to compute different types of orbits and the
associated visibility time
• Learn how to compute the reentry time

2
• Phases of a Space Mission (1/2)
– Phase A: Feasibility study (8 to 12 months)
– Phase B: Detailed Definition (12 to 18 months)
– Phase C/D: Development, Manufacture,
Integration/Test (3 to 5 years)
– Phase E: Launch Campaign
• Pre‐launch: prep. ignition, separation umbilical cables
• Launch: ignition, continuation, separation of different
phases
• Orbit transfer: launch orbit to operational orbit
• Phases of a Space Mission (2/2)
– Phase F: Mission Operations
• Commissioning phase: pull payload out, verification
• Mission operations: attitude, repositioning, orbital
maneuvers, fuel limitation  satellite lifetime
• Decommissioning phase:
– GEO: impulse to higher altitude orbit
– LEO: controlled re‐entry

3
• Necessary expertises
– Propulsion, launch and re‐entry
– Orbital analysis, navigation, attitude control and maneuvers
– Structures and materials
– Thermal and radiation control
– Sensor development
– Telecommunications
– Signal Processing
– Mechanisms
– Electromagnetic Compatibility
– Production and energy storage
– Simulation
– Project management, data distribution, etc.
• A space mission consists of:
– Launcher: transports satellite from ground to space
– Space segment: spacecraft
– Ground segment: ground station and mission control
SMOS launch
[credits ESA]

4
SMOS launch
[credits ESA]
SMOS launch
[credits ESA]

5
SMOS launch
[credits ESA]
Orbit selection depends on application and payload
Examples:
Meteosat
(Geostationary)
‐ Always looking at
the same point
‐ 3 satellites
required for
global coverage
GPS
(Medium Earth Orbit)
‐ At least 4 satellites in
view are required
OKO URSS Spy Satellite
(Molniya orb)
‐ Large fraction of the
orbital period over the
region of interest
‐ Telecom at high
latitudes and spy

6
Kepler’ s laws:
1.The secondary body describes an
elliptical orbit around the primary body.
2. The area swept by the radio‐vector going
from the primary body to the secondary
body is the same in equal time intervals.
3. The orbital period is determined by the
average distance that separates them and by
the mass of the primary
Eccentricity e Type of Orbit
0 Circular
0 < e < 1 Eliptical
1 Circular
1 < e Hyperbolic
Types of orbits:
‐ By height:
‐ LEO: Low Earth Orbit
‐ MEO: Medium Earth Orbit
‐ GEO: Geosynchronous Earth Orbit
‐ HEO: Highly Elliptical Orbit
(e.g. Tundra, Molniya)
‐ Non‐geocentric orbits: interplanetary navigation

7
6 Orbital Parameters (derived from the 2‐body Newton equation):
‐ Ω: longitude of the ascending node on the equatorial plane
‐ Ψ: inclination of the orbital plane with respect to the
equatorial plane
‐ γ: argument of the perigee at the ascending node
‐ a: semi‐major axis of the ellipse
‐ e: orbit eccentricity
‐ tp: time of passage at the perigee
(reference initial time)
 ORBIT
< 90 direct
= 90 polar
> 90 retrograde
2
2 2
2
ˆ
ˆ 0
1
( )
2
d r
r
dt r
m
mV GM
r

  
 
  
2
1 cos
h
r
e




6 Orbital Parameters (derived from the 2‐body Newton equation):
2
2 2
2
ˆ
ˆ 0
1
( )
2
d r
r
dt r
m
mV GM
r

  
 
  
2
1 cos
h
r
e




: true anomaly
e: eccentricity
0  e  1: elliptic & circular orbits
3
2 1
2 sin tan tan
2 2 1
a E e
T M t E e E M
T e
 



         


8
Orbital period
2
,
2
84,4 , 6366 ,
k GM 3,986 10 m /s
For a circular orbit:
2
,
Ex.: TRANSIT Satellite
h=1075km → v=7.3 km/s and T=106 min
Perturbations that modify orbital parameters
Osculating parameters: 6 orbital parameters as a function of time
• Asymmetry in the Earth’ s gravity field:
Earth’ s gravitational potencial
‐ ,  : Geocentric latitude and longitude
‐ Pn: Legendre polinomials
‐ Pn,m: Legendre functions (1st kind)
‐ Jn, Cn,m, Sn,m: experimentally determined coefficients
 
      
2
, , ,
2 0
1 sin
cos sin sin
n
n n
n
n
n m n m n m
n m
R
V J P
r r
R
C m S m P
r


  


 
 
  
      
 
 
    
  

 

9
Gravity Field Map (GOCE Mission)
http://www.esa.int/esaLP/ESAYEK1VMOC_LPgoce_1.html#subhead2
Orbital Perturbations (i):
‐ Atmospheric losses:
‐ Due to atmospheric drag orbit circularizes
‐ For a circular orbit:
‐ Solar radiation pressure:
‐
‐ For a perfect absorber
4.7 10 / on Earth’svicinity
‐ For a perfect reflector
2
‐ Other perturbations
‐ Moon and Sun gravity fields, tides, Earth’s magnetic field
T orbital period
M satellite’s mass
atmospheric density
R orbit radius
S projected satellite’s surface
drag coefficient (typical 2.5)
∆ 3 ρ

10
Orbital perturbations (ii):
‐ Accelerations caused by the main perturbation sources
Comparsions of the disturbing accelerations for the main sources of perturbation Forteskue et al.,
Orbital Perturbations (iii):
‐ Effects
‐ Oscillation of the orbital plane vs. nominal inclination ( Δ )
Δ
e
2
J
J
cos
‐ Regression of the nodal line
∆Ω
3
1
1
‐ Advance of the Perigee
∆ 3
2
1
1
4 5 sin
Inclination= 63.4 → zero precession (Molniya orbit)

11
LEO orbit:
‐ Polar LEO orbit: Global coverage (passes by the poles), typical
altitude between 600 and 800 km
‐ Earth‐synchronous orbit: the ground track repeats its trace over
the Earth at regular intervals
The longitude shift ∆Ω = ∆Ω1 + ∆Ω2
in one orbit for the Ecuator is due to:
‐ The Earth’s rotation (dominant term)
∆Ω 2 [rad/orbit]
Regression of the ascending node
∆Ω 3
[rad/orbit]
Earth’s rotation period:
Te = 86164,09055+0,015 ts [s]
ts: centuries from 1900
Te ~ 23h 56’ in relation to stars (24 h with respect to the Sun)
An Earth‐synchronous orbit satisfies:
n |ΔΩ| = m ∙ 2π , where n is the number of orbits
m is the number of Earth revolutions (days)

12
Sun‐synchronous orbit: the orbital plane revolves at the
same angular rate as the Earth revolves about the Sun:
1 [rev/year] ~ 1 [/day]
∆Ω 2 [rad/orbit] = 2π [rad/year]
Tes = 3,155815 ∙ 107 s (Earth‐Sun orbital period)
The satellite flies over the territory at the same local time:
‐ 1 ascending pass: the satellite flies from South to North
‐ 1 descending pass: the satellite flies from North to South
Earth‐Sun synchronous orbit:
Combination of the two previous requirements
n |ΔΩ| = m∙2π
n 2 2 2 →
1 1
∆Ω 2 [rad/orbit]
The orbit is indicated with indices n:m
∆Ω 2 (towards the West)
Example:
For a LEO: h ~550 – 950Km, T ~ 95‐100 min
→ ΔΩ~ 0,43 rad/orb
→ 2800 km separation between ground‐tracks at the equator

13
Zero Drift orbit:
‐ if m=1, there are n orbits that repeat every day:
N:m hideal (km)
14:1 894
15:1 567
16:1 275
‐ The separation between ground‐tracks is:
; ~6378
‐ Usually, this separation is excessive for Earth observation:
14:1 → d=2862,43 km
Example: LANDSAT 1,2
T=103,3 min 251:18 orbit → 251 orbits in 18 days
(almost a 14:1 orbit: 251=14∙18‐1)
Ψ=99 separation between tracks about 160 km
920

14
‐ Ground‐track examples
43:3 501:35
‐ Visibility from the tracking ground station on the Earth:
Limited visibility + limited bandwidth for downlink
limit maximum amount of data that can be downloaded

15
‐ Artemis data relay satellite:
http://spaceflightnow.com/news/n0303/20artemis/envisatartemis.jpg
• Geostationary orbit:
Coverage with 1 satellite Global coverage with 3
geostationary satellites

16
‐ If orbit inclination is not exactly 0,
inclination > 0 ‐> regression of nodal line (longitude decreases)
inclination < 0 ‐> progression of nodal line (longitude increases)
‐ This results in an “8”‐like ground track (red line below)
1
arctan
cos( ) 4


  
2

17
• HEO orbits:
Molniya orbit parameters:
a = 26560 Km (T = 12 h)
e = 0.722 (hp = 1000 Km, ha=39360Km)
=270 (Southern hemisphere perigee)
 arbitrary (depends on required coverage)
Other orbits: Solution of Newton’s equations
3 body problem: 2 rotating bodies around a third massive one
m
M2
M1
r1 r2
r
baricenter
L1
L4
L5
L2L3
Lagrangian or Libration points: equilibrium points
L1‐3 unstable; L4‐5 stable
Map the sky (L2), monitor the Sun (L1) …

18
Reentry
• Reentry depends basically on:
 atmospheric density: as a function of orbital height, solar activity… The
larger the density, the larger the friction, the sooner the reentry
 ballistic coefficient:
where M: mass, A: section, and Cd: drag coefficient.
The larger the ballistic coefficient, the later the reentry
• Recomendation: satellites reentry < 25 years
• Reentry calculations:
 NASA DAS (Debris Assessment Software):
https://www.orbitaldebris.jsc.nasa.gov/mitigation/das.html
 ESA DRAMA (Debris Risk Assessment and Mitigation Analysis) and
MASTER (Meteoroid and Space Debris Terrestrial Environment Reference)
https://sdup.esoc.esa.int/web/csdtf/home
Sample reentries computed with DAS rev. 2.02 for a 6U cubesat
h = 450 km
reentry < 2 years
h = 550 km
reentry ~ 7 years

19
• Appendix: Defining the orbit – NORAD TLE parameters
NORAD Two‐Line Element Set Format
Data for each satellite consists of three lines in the following format:
AAAAAAAAAAAAAAAAAAAAAAAA
1 NNNNNU NNNNNAAA NNNNN.NNNNNNNN +.NNNNNNNN +NNNNN‐N +NNNNN‐N N NNNNN
2 NNNNN NNN.NNNN NNN.NNNN NNNNNNN NNN.NNNN NNN.NNNN NN.NNNNNNNNNNNNNN
Line 0 is a twenty‐four character name (to be consistent with the name length in the NORAD SATCAT).
Lines 1 and 2 are the standard Two‐Line Orbital Element Set Format identical to that used by NORAD and NASA. The
format description is: Line 1
Column Description
01 Line Number of Element Data
03‐07 Satellite Number
08 Classification (U=Unclassified)
10‐11 International Designator (Last two digits of launch year)
12‐14 International Designator (Launch number of the year)
15‐17 International Designator (Piece of the launch)
19‐20 Epoch Year (Last two digits of year)
21‐32 Epoch (Day of the year and fractional portion of the day)
34‐43 First Time Derivative of the Mean Motion
45‐52 Second Time Derivative of Mean Motion (decimal point assumed)
54‐61 BSTAR drag term (decimal point assumed)
63 Ephemeris type
65‐68 Element number
69
Checksum (Modulo 10)
(Letters, blanks, periods, plus signs = 0; minus signs = 1)
Line 2
Column Description
01 Line Number of Element Data
03‐07 Satellite Number
09‐16 Inclination [Degrees]
18‐25 Right Ascension of the Ascending Node [Degrees]
27‐33 Eccentricity (decimal point assumed)
35‐42 Argument of Perigee [Degrees]
44‐51 Mean Anomaly [Degrees]
53‐63 Mean Motion [Revs per day]
64‐68 Revolution number at epoch [Revs]
69 Checksum (Modulo 10)
All other columns are blank or fixed.
Example:
NOAA 14
1 23455U 94089A 97320.90946019 .00000140 00000‐0 10191‐3 0 2621
2 23455 99.0090 272.6745 0008546 223.1686 136.8816 14.11711747148495
https://www.celestrak.com/NORAD/elements/
• Bibliography
Spacecraft System Engineering, P. Fortescue, J. Stark, and G.
Swinnerd, 2003
http://www.satflare.com/track.asp#TOP

1
T2. Space Environment and Thermal Control
• Objectives of the lesson
– Understand the Radiation in the Space Environment and how
it affects the Satellite Subsystems
– Understanding heat modeling
• Black‐body and Sun radiation
– How to design satellites that
• Avoid extreme temperatures
• Avoid large temperature changes
• Allow heat exchange

2
Space Environment
• Sun dominates the space environment of the Solar System
• It is a thermonuclear fusion reactor
• Surface is at ~5900 K
• Solar atmosphere consists of
‐ lower region extending to some thousand kilometers
Temperature peaks at 10000 K
Emits mainly UV rays
‐ upper region extending to several solar radii
Temperature is around 2∙106 K,
but very variable X‐ray emission
Space Environment
Its emission spectrum approximates to a 5900 K black body

3
Solar Radiation: Sunspots and Variability
• Sunspots are regions which are cooler than the surrounding surface.
• The larger the number of sunspots, the larger the emission of
radiation associated with solar flares: radio, X‐rays, γ‐rays
• ~ 11 year cycle
Solar Radiation: Effects on Materials
• Polymers are particularly sensitive to high energy photons
• Modifications in resistivity
• Optical changes
• Solar arrays are particularly sensitive to UV, coverglass and adhesive
subject to darkening
• May enhance erosion by atomic oxygen etc.

4
Solar Wind
• Solar Wind dominates space weather
• Consists of a flux of charged particles at 200‐800 km/s
• Earth is protected by its own magnetic field that deflects it
 solar wind shapes it.
• Particles trapped in the Earth's magnetic field “go” to the Van Allen
belts.
• Only observable on Earth when it is strong enough to produce boreal
aurorae, geomagnetic storms or plasma tails in comets.
Van Allen Radiation Belts
• Contain energetic particles trapped in the Earth’s magnetic field
• Inner belt (100‐10000 km) has a larger concentration of protons
• Outter belt (13000‐60000 km) consists mainly of electrons
• The effects caused by these particles are
degradation of electronics due to accumulated radiation dose
degradation of solar array performance
single event upsets (SEUs)
dielectric discharging

5
Relative radiation level experienced in different orbits
(from STMicroelectronics)
Calculation can be done using software tolos such as:
 SPENVIS (SPace ENVironment Information System)
https://www.spenvis.oma.be/
South Atlantic Anomaly:
the “Bermuda triangle of the spacecrafts”
• Region where the inner Van Allen belt is closer to the Earth
• Exposes satellites flying below the belt to higher radiation doses,
causing disruptions on systems.
• Caused by the non‐concentrity of Earth’s magnetic field.

6
Geomagnetic Storms
Disturbance of Earth’s magnetosphere as a consequence of solar ions caused
by disruption
Effects:
Satellite systems: Particle damage, UV, Atmospheric drag
Power Systems: Geomagnetically induced currents
Navigation Systems: Changes in ionosphere
Communication: Ionospheric irregularities
Single Event Effect
A Single Event Effect is caused by ions or electro‐magnetic radiation
striking a sensitive micro‐electronic device.
• Single event upset (SEU): A change of state or transient
• Sigle event Latchup (SEL): Event causes loss of device funcionality
• Sigle event Burnout (SEB): Device destruction is caused due to a hight
current state in a power transistor.

7
Examination of nearly 1300 single‐event upsets from one
computer on the TAOS mission shows that nearly 50 percent
occurred in the South Atlantic Anomaly, whereas only 5 percent
of orbital time was spent there.
Thermal Control
- Extreme Environment
- Space ~ 3 K
- Intense solar radiation
- Eclipses
- Vacuum: no air circulation ⇒ risk of overheating
- Narrow temperature ranges for devices
- 0 to 50C for batteries
- ‐10 to 150C for most electronics
- 7 to 35C for fuel
- Large and fast thermal gradients
- Alignment distortion
- Fast expansion and contraction
- Risk of material cracking

8
Radiation
– Sun emits approximately as a black body
– A black body is an ideal body that…
absorbs all the energy at all wavelengths from all directions and
polarizations, and in thermal equilibrium re‐emits it all following
Plank’s law
‐ Radiation equations
1. Wien’s law: energy peak at

σ= Wien’s displacement constant (~2898 μm∙K)
2. Planck’s law: energy density by wavelength
,
2 1
1
3. Stephan‐Boltzmann’s law (will be seen later)
[http://www.oswego.edu/~kanbur/a10
0/lecture13.html]
• Radiation
– Sun temperature: ~ 6000K Earth temperature: ~ 300K
⇒ λmax= 483 nm (visible, blue) ⇒ λmax= 10 µm (thermal infrared)
[http://scienceofdoom.files.wordpress.com/2009/11/blackbody_curve‐sun‐earth.jpg]

9
• Radiation
– A black body is ideal, the Sun is not
– Real bodies have
• Absortivity: α
– fraction of energy absorbed, relative to that of a black body
• Emissivity: ε
– fraction of energy emitted, relative to that of a black body
– Atmosphere also modifies spectrum
Black‐body (blue) / Sun TOA (red)
• Radiation reaching the satellite
[Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003]
Typical Spacecraft Thermal Environment

10
• Radiation reaching the satellite
– Direct light from the Sun ~1400 W/m2
– Albedo (reflected light from the Sun on the Earth’s surface) ~30%
of direct solar radiation
– ~10‐40% over soil
– ~5% over water
– ~40‐80% over clouds
– Earth’s radiation ~250 W/m2
• Internal heat dissipation
• Radiation emitted by the satellite
– Because of body emission
Etotal=σ∙T4 [W/m2] (Stephan Boltzmann’s Law)
[https://engineerjau.wordpress.com/2013/06/09/where-does-the-stefan-boltzmann-constant-come-part-1-of-2/]
• Heat balance
– Equilibrium: Qin = Qout
– Body illuminated by the Sun:
Pdirect + Palbedo + Peartshine + Pdissipated = Pemitted
• For GEO orbit only direct light is significant
– Body in shadow:
Peartshine + Pdissipated = Pemitted
– Depends on absorptivity and emissivity

11
• Heat balance
– Absorbed
Pdirect =Direct solar∙Silluminated∙αs [W/m2]
Palbedo=Albedo ∙Silluminated∙αs [W/m2]
Pearthshine=Earthshine ∙Silluminated∙ε [W/m2]
Pdissipated=Power [W]
– Emitted
Pemitted = Stotal∙ε∙T4∙σ
• Absorptivity and emissivity examples
– White paint: α=0.15,   ε=0.9
– Black paint: α=0.9,   ε=0.85
– Aluminum: α=0.15,   ε=0.05
– Gold: α=0.25,   ε=0.04
– Solar cells (Si): α=0.75,   ε=0.82
– Solar cells (AsGa): α=0.88,   ε=0.80
• Heat balance dynamics
C: heat capacity

Transient‐state
Oscillation in steady‐state

12
• Example 1:
– Orbit time 94 min, 63 min in light
– Considering 1 kg of aluminum, 10 cm each side
– 1 day simulation
– Materials: Silver (α=0.37,  ε=0.44) ⟶ blue
AsGa (α=0.88,  ε=0.80) ⟶ red
Si (α=0.75,  ε=0.82) ⟶ green
Silver (blue, ‐7 to 3 °C), AsGa (red, ‐12 to 10 °C), Si (green, ‐17 to 3 °C)
[°C]
Time [∙104 s]
• Example 2
– Orbit time 94 min, 63 min in light
– Considering 1 kg of aluminum, 10 cm each side
– 1 day simulation
– Materials: Gold (α=0.25,  ε=0.04)
Gold: 167 to 175 °C
[°C]
Time [∙104 s]

13
• Thermal design (1/3)
– Thermal control requires a balance of onboard heat emission and
absorption
– Radiation surfaces have to be sized large enough to keep temperature
within upper temperature limit during solar exposition
– Whole system has to be designed respecting security margins
– Coated and painted surfaces have to be designed according to valid
temperature ranges
– Inner sides are black painted to maintain homogeneous temperature
– Heaters have to be sized large enough to keep temperature above the
lower temperature limit during umbra periods
– Non‐radiating surfaces are generally insulated with multilayer
insulation
– It is highly recommended to simulate internal temperature gradient
(Thermal Desktop, ESATAN…)
• Thermal Systems
– Passive systems
• No power consumption
• High reliability
• Low heat transfer capabilities (except heat pipes)
– Active systems
• Power and telemetry requirements
• Lifetime considerations, operational risk
• Adaptable, flexible and higher heat transfer capabilities

14
• External Systems
– Passive systems
• Radiator surfaces
• Coating (paint)
[http://novitastechnology.com/Projects.html]
– Passive systems
• Thermal Blankets
MLI: Multi‐Layer Insulation
[http://en.wikipedia.org/wiki/Multi‐layer_insulation]

15
– Passive systems
• Sun Shields
http://www.jpl.nasa.gov/images/aquarius/aquarius‐20090429‐browse.jpg
– Active systems
• Louvers

16
• Internal Systems
– Passive systems
• Heat Pipes (high transfer capability)
• Radiating systems
[http://hectorleonosorio.blogspot.com/2010/07/tubo‐
de‐calor‐heat‐pipe.html]
http://en.wikipedia.org/wiki/File:Laptop_Heat_Pipe.JPG http://www.cheresources.com/htpipes.shtml
• Internal Systems
– Active systems
• Heaters
• Peltier cells
[http://mechanical‐
engineering.esa.int/thermal/images/heater3big.jpg]
[http://en.wikipedia.org/wiki/File:Peltier_(detail)_LMB.png]

17
• Design Verification
– Similarity
• Use of space qualified materials
– Verification Analysis
• Comprehensive Thermal Math. Model (CTTM)
– Thermal Vacuum Test
• Temperature ranges
• Temperature gradients
• Validation of CTTM
• Example: SMOS at Large Space Simulator

18
• Bibliography
o Spacecraft System Engineering, P. Fortescue, J. Stark, and G.
Swinnerd, 2003
o ECSS‐E‐ST‐10‐04C
Space Environment (15 November 2008) in Space Engineering
Method for the calculation of radiation received and its
effects, and a policy for design margins (15 November 2008)
in Space Engineering
Spacecraft charging (31 July 2008) in Space Engineering

6/26/2018
1
T3. On Board Data Handling
IEEE Young Professionals in Space, Barcelona July 17-21 2018 - © UPC-BarcelonaTech - 2018 1
Objectives of the lesson:
‐ Understand what “On Board Data Handling” is
‐ Understand what kind of telemetry data is required
‐ Architecture of the OBDH
‐ Issues affecting the OBDH design

6/26/2018
2
Introduction
• Involved subsystems.
• Types of data, requests, actions….
• Flight software and On‐Board Computer: architecture, OS, CPU.
Payload
Recording
device
EPS, ADCS…
Communications Subsystem
On-Board Computer
TT&C
Ground
segment
Payload data
Configuration
Logs,configuration
Firmwareupdates
Mission control, Sw. updates
Digital buses
Introduction (cont’d)

6/26/2018
3
Telemetry data (downloaded to GS):
• Housekeeping data:
 Monitor the status of the s/c.
 Sensors, device states, system
logs (e.g. error messages).
• Attitude data:
 Determine/monitor attitude of
s/c.
 Mode of operation, controller
state, sensor readings.
• Payload data:
 Payload data is needed to
monitor the status of the
payload and the experiments to
be processed before being
transmitted to the ground.
Spacecraft operations
Telecommands (uploaded to s/c):
• Control modes of operation:
 Perform initial commissioning.
 Enable states in spacecraft FSM.
 Handle contingency situations.
 Change payload modes or setup.
• Configure devices and subsystems:
 Change parameter values (they
have been designed to allow that).
• Fix errors (to some extent):
 Upgrade software and firmware.
 Access digital buses and
microcontrollers.
Telemetry data – housekeeping data
• Data volume is critical (constrained bandwidth, data‐rate and
visibility).
• Number of different parameters to monitor.
• Low sampling frequency (period from 1 s to 2min)
• Telemetry data can be transmitted to ground
• in real time (e.g. beacon signals)
• stored (delayed transmission in case of link unavailability)
THE 3CAT‐1 CASE:
9600 bps, MTU: ~64 kB, Selective repeat ARQ protocol
→ 1 min (best case, no errors in protocol/handshakes.)

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4
Telemetry data – housekeeping data
• Data volume is critical (constrained bandwidth, data‐rate and
visibility).
• Number of different parameters to monitor.
• Voltages, Current Consumption, ADCS…
• Low sampling frequency (period from 1 s to 30 s)
• Telemetry data can be transmitted to ground
• in real time (e.g. beacon signals):
• Very useful for amateur bands,
• stored (delayed transmission in case of link unavailability):
• Just for critical information that cannot be lost.
THE 3CAT‐2 CASE:
9600 bps, MTU: ~26 kB file chunks, Selective Repeat ARQ file oriented
protocol  45sec/chunk (medium case, assuming 5% of PER, packet
size=125 bytes)
Telemetry data – housekeeping (example)
Type Source Analog / Digital
Temperature
Solar panels, batteries,
OBC
Analog (thermistors,
RTD, lcs,…)
Voltages & currents
Power supplies solar
panels, power buses
Analog (sensing
circuitry)
Operation modes
Payload status, heaters
(on/off)
Digital
Actuation / deployment
mechanisms
Solar panels and
antennas deployment
Digital

6/26/2018
5
Telemetry data – housekeeping data, considerations
• Number of required acquisition channels
• Accuracy required
• Precision of any telemetry channel shall allow an assessment of
the nominal or out‐of‐tolerance status of the monitored
parameters
• Sufficient telemetry shall be available to allow the verification of:
• Power budget with accuracy better than e.g. ±5%
• Temperature with an accuracy better than e.g. ±1%
• Types of sources.
• Format and storage (standardization is critical):
• e.g. Analog voltage 0 to 5 VDC: units are [V] or [mV]?
• Storage type: 8/16/32‐bit integer, floating point, ASCII (“5123” =
5.123 V)
THE 3CAT‐1 CASE: EPS HOUSEKEEPING
Data volume without overheads ~63.3 kB = ~8000 registers (timestamp + data).
5 sensing magnitudes (up to 7 channels) @ Ts = 1 min. → < 5 hours of history.
Telemetry data – housekeeping data, considerations
• Number of required acquisition channels
THE 3CAT‐2 CASE: EPS HOUSEKEEPING
Three different logs ~96 kB @ ~38 bytes/register = ~2500 registers
Compressed  Down to 20kB file size (high compression ratio)
Lots sensing magnitudes at a rate of 20 seconds  ~12‐13 hours  1 log file
!!!! Which is the revisit time over the Ground Station !!!!
THE 3CAT‐2 CASE: ADCS HOUSEKEEPING
Three different logs ~96 kB @ ~84 bytes/register = ~1150 registers
Compressed  Down to 40kB file size (medium compression ratio)
Lots sensing magnitudes at a rate of 5 seconds  ~1.5 hours  1 log file
!!!! Which is the orbital period !!!!

6/26/2018
6
Telemetry data – attitude determination
• Attitude determination can be based on:
‐ Sun and Earth sensors
‐ Accelerometers
‐ Gyros
‐ Magnetometers
‐ Star trackers
• Few data is normally required
• Sampling rate
Normally low sampling rate (T = 1~ 60s) even if it can be higher on
demand
(ex.: stabilization of the spinning)
Telemetry data – payload data
• Data volume is variable. Different payloads will generate different
amounts of data.
• Ultimately limited by the link budget (not the storage).
• Sampling rate:
• May be constant and regular on time (ex.: to study a degradation
in space conditions)
• Fast sampling bursts (ex.: to study fast dynamic changes of a
MEMS experiment)
• Experiments may be disabled after their duty cycle in order to
release storage regions to other tests.
• Payload data is not necessarily kept forever in the recording device.
Downloading it may mark the recording device region as “available”
again.

6/26/2018
7
Telecommand data (example)
DESTINATION EXAMPLES
Spacecraft
Orientation (activation of wheels & gyros)
Micro‐propulsion systems
Manual activation of mechanisms (heaters)
C&DH unit
Changes of operative modes (stand‐by, energy saving,
active modes)
Data upload/download
Power section
Change on battery level based operation modes
Reset
RF Rx/Tx Beacon On/Off
Payloads Power on/off, sampling rates, activation timer, …
Telemetry data packet
• Telemetry data is encapsulated into a packet and prefaced by a packet
header
• Packet header contains:
‐ Source ID
‐ Number of the current packet in the sequence of packets produced
‐ Length of the data field
‐ Packet error control can be added.

6/26/2018
8
Telecommand data – packet telecommand layers
• Packetization layer
‐ TC packet consists of a fixed length packet header and a variable
length data field
‐ Header contains: destination ID, packet length, packet control
error features (optional)
• Segmentation layer
Combination of multiple TC packets with data contained inside
• Coding layer
Provides the error correction and synchronization required by the
spacecraft
• Physical layer
RF link itself
Telecommand data
• Elementary commands or complex sequences (SW update)
• Validated by the OBDH unit before its execution
‐ OBDH transfers the commands, normally to the OBC
‐ Internal communication protocols (I2C, SPI, RS232, …)
‐ OBDH confirmation to the ground of the data received before its
execution.
• Telecommand can be executed instantly or deferred

6/26/2018
9
Spacecraft Flight Software: general architecture.
• Common topologies: centralized or distributed.
C&DH
Subsys. 3
Subsys. 1
Subsys. 4
Subsys. 2
Subsys. 4
Integral modules (hardware)
connected through digital buses or
software modules interfacing with
OBDH core.
Subsys. 1
OS
Framew
ork
Subsys. 1
OS
Framew
ork
Subsys. 1
OS
Framew
ork
Spacecraft Bus
e.g. Voyager, Galileo (NASA’s outer
planetary spacecraft), AAUSAT3
(nano‐satellite mission)
Spacecraft Flight Software:
• Structured software that is:
• Reliable:
• Precludes error propagation from one module to others;
• Prevents deadlocks;
• Controls timeouts, watchdogs, event sequencing, transitions to safe
states;
• Detects invalid input data;
• Protected memory regions (isolation);
• …
• Automatically analyzable: verify code flow, data integrity, deadlocks,
infinite loop conditions, etc. before compile‐time! (Software verification
processes are usually hard and tedious.)
• Traceable: it should be easy for ground operators to determine what is
happening with reduced information.
• Reusable: standard interfaces, modular architectures, encapsulation of
components.
• Modifiable (at design‐time), and updateable (at run‐time!)
Specially valuable when missions have short development cycles (e.g. nano‐satellites)

6/26/2018
10
• Implementation of multiple‐level Finite State Machines.
• High‐level mission states → internal sub‐states → subsystem modes
→ device commands.
• Should always guarantee spacecraft safety, i.e. tend to be complex
and should be automatically analyzable.
• E.g. if contingency: SYSTEM RESET (may take from several seconds
to some hours), SAFE STATE (only critical subsystems are enabled).
• E.g. 3Cat‐1 energy and system FSM (each high‐level state is
decomposed into several routines):
CONTINGE
NCY
STANDBY
PAYLOADS
COMMS
IDLE
INIT
LE ∨ IE
LE ∨ S
LE ∨ S
Controlled
by EPS
OBC shut
down
Energy
critically
low
OBC
shut
down
Low-power mode
Comms RX enabled
Payloads disabled
Nominal power
mode
Comms. full
mode (RX & TX)
Payloads
enabled
0% 10% 15% 75% 100%
OBC
power-off
OBC
power-on
State of
charge
• On‐board pre‐processing of data: only valuable data is downloaded.
• Failure Detection, Isolation and Recovery techniques, at software
level. Systems that are capable of identifying or recovering from
errors without orders from ground operators:
• E.g. when round‐trip delay times are too long to wait for the
ground operators to react.
• Current trends: autonomous spacecraft.
• “Minimum” human intervention: on‐board Mission Planning and
Scheduling systems.
• On‐board identification of interesting data acquisition
opportunities.

6/26/2018
11
• Real‐time operating systems (RTOS):
• Deterministic latencies RT scheduler:
predictable execution.
• RT ≠ fast compu ng.
• Priority‐based. Preemptive / non‐preemptive.
• Programming aspects: kernel manages Real‐
Time Tasks (similar to threads).
• Examples of RTOSes used in spacecraft:
(*) http://opensource.gsfc.nasa.gov/projects/cfe/
• Common frameworks and standards:
• NASA cFS/cFE (Open Source, Free License) (*). Available for RTEMS, Linux, VxWorks…
• TSP: Time and Space Partitioning (Avionics Application Standard Software Interface
ARINC 635 standard.)
Task
(NAME / PRIORITY)
A / 1
B / 2
C / 3
Execution on CPU
Task context switches are managed
by the kernel scheduler (i.e. they
are deterministic, predictable).
 VxWorks: proprietary.
 RTEMS: open source, free license.
 Micrium uC/OS‐III: open‐source.
 QNX: proprietary.
 LynxOS: proprietary.
 FreeRTOS: open‐source, free license.
 Xenomai (Linux patch): soft‐real‐time,
open‐source, free license.
 PREEMPT_RT (Linux patch): soft‐real‐
time, open‐source, free lic.
CPU and OBC boards
Commercial-of-the-shelf components are used (for CubeSats).
Pumkin FM430
http://www.cubesatkit.com
Nanomind A712C
http://gomspace.com/
Intrepid
http://tyvak.com/products/products.html
CPU MSP430, 16bit @8 MHz ARM7, @ 8-40 MHz ARM9, 32bit @ 400 MHz
Memory 55 kB Flash, 5 kB RAM 2MB RAM, 4 MB Code
storage, 4 MB data flash
64 MB SDRAM, 128 MB NAND Flash,
128 byte EEPROM
CPU Power
consumption
From 0.2 μA to 330 μA
(@1 MHz)
3.3 V, 70 mA (nominal) From 0.9 μA to 80 mA
OS SalvoOS, FreeRTOS FreeRTOS OS Linux 2.6.x, OpenEmbedded
Integrated
peripherals
Direct Memory Access
controller, 12bit ADC and
DAC, 2 Timers with pulse
with modulation, 2USART
with SPI, I2C or
Asynchronous UART, 6
8bits I/O ports
CAN and I2C interfaces
RTC - real time clock
On-board magnetometer, 3x
PWM drivers, 6x sun-sensor
inputs, 3x Rate-gyro inputs…
10/100 Ethernet, USB2.0, 2 USB Host, 1
USB Device, 1 Synchronous Serial
Controller, 1 SPI, I2C, 64 digital I/O ports,
16bit parallel bus, 5 serial interfaces
(UART/USART), External SD Card
interface, 4 10-bit ADC

6/26/2018
12
Issues affecting OBDH (i)
• Radiation:
• Total dose: for a unit installed externally of the ISS a total dose of 3 kRad
for a 3‐years mission is foreseen with a box thickness of 3 mm Al
equivalent, safety margins applied
• Total dose mainly due to X‐rays and Gamma‐rays.
• Components sensitive to total dose present functionality anomalies
(power consumption abnormal increase, ….)
• Single event effects:
• SEE: generalized category of anomalies resulting from a single ionizing
particle
• SEU: Single event upset (non‐destructive ), SEL single event latchup
(destructive), SEGR single event gate rupture (destructive), …
• Radiation effects on design:
• Radiation limits the number and the type of components available to the
designer : uController, CPUs, FPGA Field Programmable Gate Arrays ,
EEPROMs, Amplifiers, Voltage regulators, etc
Issues affecting OBDH (ii)
• Radiation: solutions
• Radiation‐hardened devices:
• Tolerate higher radiation doses.
• Special manufacturing techniques.
• Bipolar technology (not CMOS)
• Shield components.
• On CPU: usually SPARC or PowerPC architectures.
• E.g. SPARC‐V8 LEON3 (VHDL‐synthezisable, free license),
project started by ESA.
• Error‐Correcting Codes, e.g. NAND‐Flash with 4‐bit BCH ECC.
• Redundancy: at device‐ or circuit‐level. (Software redundancy could
also be possible)
Block #1
Block #2
Block #3
Input data Arbitrator
(voting)
Input data
1011
1001
1011
1011
1011

6/26/2018
13
Issues affecting OBDH (iii)
• Thermal Constraints:
• Thermal environment and vacuum limits the types and number
of components available to the designer.
• Typical operational temperature range for a spacecraft OBDH
unit could be –20 °C to +50 °C
• Electronic components are usually classified and sold in three
different families:
• Commercial range: 0 to +70 °C ( plastic package)
• Industrial range : –40 to +80 °C
• Military range : –55 to +125 °C ( ceramic package)
[sometimes for automotive applications as well]
• Heat from electronic components can be removed only by
conductive paths (heat pipes, no fans, no finned heat sinks....)
or by radiation

6/26/2018
1
T4. Attitude Determination and
Control System
T4. Attitude Determination
and Control System
Control System
Objetives of the lesson:
‐ Payload and sensors requirements for operation (e.g. Telescope)
‐ Mission operation pointing requirements
‐ Power demand: Sun‐pointing
‐ Communications: antennas pointing
‐ Thermal regulation: avoid heat gradients, heat dissipations to deep
space
‐ Thruster pointing for orbital change
‐ Mission stat

6/26/2018
2
Control System
Basics: Linear Momentum
• Center of mass M∙roc = Σ m∙rOP
Lets us to handle withe the full body as an equivalent particle (e.p)
of mass M in the position roc. In that position Σ(m∙rcp) = 0 is
acomplished always.
• Linear momentum L = M∙v
• Basis to translation/orbit dynamics
• Impulse L
0
• An impulse is equal to the change in the momentum that it changes
Control System
Angular Momentum
Moments of inertia: Ixx, Iyy, Izz e.g. Ixx = (y² + z²)∙dm
Measures the distance from an axis: (y² + z²) = d² from x‐axis
Products of inertia: Ipq e.g. Ixy = xy∙dm
Matrix of inertia
As an equivalent to the total mass of the body, it relates the
angular momentum with the angular speed vector

6/26/2018
3
Control System
Angular Momentum
Torque Impulse T
0
Equivalent to the linear momentum, but since it is not just
proportional to the total mass, but it depends on a matrix with
products of inertia acting as cross‐coupling among diferent axes, it
is not as easy understandable.
The form:
ω] = [Ic]∙T ,
helps to understand it.
Control System
Momentum Build‐up
• As in the linear case, only external torques affect the total angular
moment of the system because
Σ Text = 0 → H is constant
• External torques could be either on board torquers, such as thrusters
or magnetorquer rods, or external disturbances (eg. Solar wind).
• Whether the spacecraft has a momentum or not, Attitude Control
System (ACS) has to manage the momentum in order to get the desired
performance or to damp the external disturbances

6/26/2018
4
Control System
Attitude motion of specific types of spacecraft
Spacecraft
Without momentum bias (MB) With momentum bias
3 axis stabilized (no MB) Spinner Hybrid
Partially de‐spun 3 axis stabilized
(with MB)
Control System
Momentum bias
As in the linear case with L, momentum reduces ther
sensitivity to torque → gyroscopic rigidity.
ψ
Momentum bias is used by spinning‐up the spacecraft before
rocket firing to avoid misallign with course. Then, it is usually
spun down.
When a torque is applied to a biased body, it introduces a
nutation mode. ACS must damp it. When the torque stops, it
introduces more nutation oscillation, a good timing has to be
set in order to compensate the new one with the former.

6/26/2018
5
Control System
Spacecraft → With momentum bias → Spinner
METEOSAT‐1 METEOSAT‐3
Font: http://es.wikipedia.org/wiki/Archivo:METEOSAT.gif
MSG‐3 is the third in a planned series of four satellites operated by
Eumetsat, the European Organisation for the Exploitation of
Meteorological Satellites
Read more at: https://phys.org/news/2012‐03‐msg‐satellite‐ready‐
weather‐monitoring.html#jCp
Control System
Simple spinner
Payload and other subsystems are placed in the spinning section
• Power: Solar array limited to sides
• Thermal: avoids heat gradients
• Communications: non directive anthennas
• Limited mounting space for non‐scanning payloads
• Spin Axis: Least intertia axis could only be chosen for a limited
duration (rockets). For a satellite demanding a preservation of the
spinning attitude, maximum inertia axis must be chosen.
• Nutation: spinning introduces an oscillatory mode.
 requires the use of passive dampers or control torquers

6/26/2018
6
Control System
[https://www.youtube.com/watch?v=R8jiU1EYIAM]
Spacecraft
With momentum bias → Spinner → Partially de‐spun
GIOTTO
Control System
Dual spinner
Payload, antenna and feed are placed on a de‐spun platform
• Power and thermal control similar to the simple spinner
• Spinning parts inertial axes should be balanced
• As a spinning body, nutation occurs and has to be controlled by ACS

6/26/2018
7
Control System
Spacecraft
With no momentum bias (no MB) → 3‐axis stabilized
Control System
Image font: http://www.esa.int
Spacecraft
With no momentum bias (no MB) → 3‐axis stabilized
e.g. SMOS
(and most other
Earth Observation
satellites)

6/26/2018
8
Control System
Image Font: http://www.ask.com
/wiki/Magellan_probe
Dual spinner
• Instead of spinning the spacecraft, wheels allow to
keep it stabilize it.
• Wheels store momentum and changes satellite
facing by varying its speed.
• 3‐axis wheel stabilization allows control almost
independently for each axis by placing them
ortogonally.
• Usually, a 4th wheel is added in a combination
affecting all of the axis in order to have redundancy
and reliability.
• Momentum wheels accomodate fluctuations of 10% of the bias
Control System
ACS Block Diagram (i)
‐ Attitude and Control System keeps the spacecraft
in a given attitude.
‐ Compensates all possible disturbances
appropriately.
‐ Furthermore, it has to let the spacecraft perform the attitude changes for
the correct developement of the mission. Typical example is to set the
pointing to the Earth or the Sun.
‐ Desired torques combined with disturbances affect the spacecraft and are
sensed.
‐ On‐board and on‐ground computers compute the torque to be applied to
correct and/or change the attitude following the established plan.
Image font: Spacecraft Systems Engineerig. Fortescue, Stark, Swinerd

6/26/2018
9
Control System
ACS Block Diagram (ii)
Control loop:
PID control is by far the most common
Proportional: proportional to error
good response to quick disturbances
Integrator: proportional to the accumulated (“integrated”) error
good to remove biases
Derivative: proportional to rate of change of error
good for stability
Pointing and momentum management are usually controlled independently
Control System
External Torques
Disturbances
• Aerodynamic < ~500 km
• Magnetic ~ 500 – 35000 km
• Gravity gradient ~ 500 – 35000 km
• Solar radiation > ~ 600 km
Controllables
• Gas Jet
• Magnetorquers
• Adjustable geometry

6/26/2018
10
Control System
Gas Jets
• Requires fuel
• Suitable for any torque size
• ON‐OFF control only
Control System
Magnetorquers
• No fuel, but power required
• Complicated Earth field
• No torque about field line
https://www.cubesatshop.com/product/isis‐magnetorquer‐board/

6/26/2018
11
Control System
Adjustable geometry
• Require fuel
• Suitable for any torque size
• ON‐OFF control only
Control System
Internal Torques
- Internal Disturbance Torques
- Mechanisms (deploying parts)
- Fuel movement
- Controllable Internal Torquers
- Reaction wheels
- Momentum wheels

6/26/2018
12
Control System
Attitude Sensors
- In order to keep a good attitude, attitude must be measured.
- Reference sensors, with external information sources
- Stars (1 arcsec)
- Sun (1 arcmin)
- Earth (horizon) (6 arcmin)
- Magnetometer (30 arcmin)
- GPS (6 arcmin)
- Inertial sensors
Control System
Attitude Sensors
‐ Reference sensors, with external information sources
‐ Stars (1 arcsec)
‐ Fixed‐head star tracker: electronically scanned imatge
. CRT technology
. High power and mass
. Magnetic interference

6/26/2018
13
Control System
Attitude Sensors
. Reference sensors, with external information sources
. Stars (1 arcsec)
. Star scanner
‐ No moving parts
‐ Mounted on spinning spacecraft
. CCD Star tracker (and mapper)
https://www.cubesatshop.com/product/mai‐ss‐space‐sextant/
https://www.cubesatshop.com/product/nst‐1‐nano‐star‐tracker/
Control System
Attitude Sensors
. Sun Sensors
. Single axis (and two axis by orthogonal combination)
. Two‐axis CCD array Sun sensor
https://www.cubesatshop.com/
product/nss‐cubesat‐sun‐
sensor/
Photodiode (1 per face)
2‐Axis digital sun sensor
https://www.cubesatshop.com/product
/ssoc‐d60‐2‐axis‐digital‐sun‐sensor/
https://www.cubesatshop.com
/product/mai‐ke‐sun‐sensor/
1‐Axis digital sun sensor

6/26/2018
14
Control System
Attitude Sensors
. Earth‐horizon
. Infra‐red spectrum sensor
. Target always present
. Slower response
https://www.cubesatshop.com/product/mai‐ses‐ir‐earth‐sensor/
Control System
Typical Satellite Attitude States and Transitions (i)
• After deployment the satellite (CubeSat) is completely disconnected for
TBD minutes in order to avoid any perturbation to the main passenger.
• Detumbling is responsible to slow‐down satellite oscillations and stabilize
it using 3‐axis magnetorquers: B‐dot algorithm.

6/26/2018
15
Control System
Typical Satellite Attitude States and Transitions (ii)
• Once the derivative of the magnetic field intensity is smaller than a
predefined value ( ∗
), it will transition to another mode if
o Nominal if the EPS battery level is above a given threshold ( ∗
), or
o Sun safe if the EPS battery level is below a given threshold ( ∗
).
• In nominal mode reaction wheels are activated and will point e.g. the
antenna boresight or the camera optics to nadir or to the target.
• However, if the above conditions are not met anymore, a return to the
“Detumbling”
• Transition from Sun Safe Mode (Sun pointing) to Nominal Mode occurs
when the EPS battery level is above a given threshold ( ∗
), and vice‐versa
when the EPS battery voltage is below a TBD % of its maximum value (TBC).
• Transition from Sun Safe Mode to Survival Mode occurs when the EPS
battery level is below a TBD % of its maximum value (TBC).
Control System
All transitions are commanded from ground
(except transitions to emergency states)
Typical Satellite Attitude States and Transitions (iii)

6/26/2018
16
Control System
Bibliography
Spacecraft System Engineering, P. Fortescue, J. Stark,
and G. Swinnerd, 2003

1
T5. Electrical Power Supply System
• Motivation
– All systems need energy to operate
– Solar energy is unlimited and cheap, but not available for all missions
– Sometimes a secondary power supply is need
• Learning objectives
– Know different energy‐supply options
Understand pros and cons of each option
– Know how to dimension the energy system
– Understand that performance degradation requires oversizing the
whole system
– Know the typical blocks of an EPS

2
• Design process
– Power System depends on
• Mission type (Earth orbit, inner planets, outer planets…)
• Mission life
• Mission payload
– Power System Block Diagram
Primary
Source
Primary
Source
Power
Management
Batteries
(Secondary
Source)
Power
Regulation
Regulated
Bus
Primary
Sources
• Power Sources
– Main
• Solar Cells
• Fuel Cells
• Radio‐isotope Thermoelectric
Generator (RTG)
• Nuclear fission systems
• Solar Heat Systems
– Secondary
• Batteries Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003
System W/kg Power Available Cost $/W
Solar Array 67 25 kW 600
R.T.G. 5 285 W 50‐60K
Solar Dynamics 25 > 20 kW ‘High’
Fuell Cell 110 12 kW 40
Spacecraft System Engineering, Fortescue, Stark, Swinerd, 2003
Power Sources Performance Comparison

3
• Solar panels
– Group of Solar cells
[http://photojournal.jpl.nasa.gov/catalog/PIA14172]v
• Solar Cell
– Is a p‐n semiconductor device that converts photons in electrons
• With no illumination, the p‐n junction achieves an equilibrated
state ⇒ no current flow
• With illumination, photons with enough energy (gap band)
create electron‐hole pairs that create a potential difference in
the cell’s terminals ⇒ current flow
http://physics‐tutor.site90.net/drupal/node/2solarcellcentral.com

4
• Solar panel components
– Gap band is the minimum energy amount needed to allow
photoelectric effect
• It depends on material properties
– Energy depends on frequency because of Plank’s law
λ
Where h is Plank’s constant (6.62∙10‐34 J∙s)
• ⇒ Each material converts photons to energy until a maximum
wavelength
Material Band Gap (eV) Maximum Wavelength (μm)
Si 1.12 1.12
CdS 1.2 1.03
InP 1.344 0.923
GaAs 1.35 0.92
CdTe 2.1 0.59
GaP 2.24 0.554
Spacecraft System Engineering, Fortescue, P., Stark, J., Swinerd, G. pag. 329
http://www.ioffe.rssi.ru/SVA/NSM/Semicond/InP/bandstr.html
• Solar Cell Electrical Model Diagram
IPH D RP
RS
‐ ‐

5
• Solar Cell Electrical Model
– Current through load is maximum when ZLoad≈0 (shortcircuit)
• Imax≈IPH
– Voltage is maximum between terminals when ZLoad ≈∞ (open circuit)
• Vmax ≈IPH∙RP
– Transferred power is maximum when load is adapted to cell
• Solar Cell P‐V & I‐V graph example: Spectrolab
Blue: I(V[Volts]) [mA]
Red: P(V[Volts]) [mW]
• Solar Cell P‐V & I‐V graph example
– We can see clearly the Maximum Power Peak (MPP)
– It depends on load impedance and incident power:
• With less incident power, output voltage and current will be lower
– It is recommended to set working point at MPP to get the maximum
power available
• Use of Maximum Peak Power Tracking (MPPT)
– Available power on solar cells depends on
• Light power density (∼1400 W/m2)
• Conversion efficiency
– Manufacturing material
– Temperature
– Absorbed radiation
• Solar cell surface

6
• Solar Cell Material Comparison
Material Si GaAs InP
Efficiency 15,5% 19,5% 16,5%
10 years GEO EOL 8,8% 13,6% 16,2%
Power loss/°C 0,438% 0,162% 0,206%
Mass density 0,55 kg/m2 1,69 kg/m2 1,77 kg/m2
Relative Cost 1 5 36
• Radiation damage in solar cells
– Radiation damages cells
• Efficiency decreases
• Less available power
[From Solar Cell Array Design Handbook by Rauschenbach, H. S.]

7
• Temperature effect on solar cells
– If temperature raises, efficiency decreases
[From Solar Cell Array Design Handbook by Rauschenbach, H. S.] [http://www.altestore.com/howto/Solar‐Power‐Residential‐Mobile‐
PV/Off‐Grid‐Solar‐Systems/Electrical‐Characteristics‐of‐Solar‐Panels‐
PV‐Modules/a87/]
• Solar array configurations:
– One single solar cell does not give enough current and voltage
for almost any application
– Solar cells can be connected in serial or parallel
• Serial: output voltage is the sum of cell voltages
• Parallel: output current is the sum of cell currents
• In both situations, output power is the sum of powers of
the different cells

8
• Solar array forming problems
– In parallel blocks, if one cell is not
illuminated, its voltage will be 0 V
 Current of other cells will flow through
that cell
 Block voltage will be 0 V (short‐circuit)
 Cell damage risk
+
‐
Illuminated
solar cell
Not
illuminated
solar cell
‐ In serial blocks, if one cell is not
illuminated, its current will be null
 Current of other cells will not flow
trough it
 Output current is null
+
‐
Illuminated solar
cell
Not illuminated
solar cell
If a solar cell fails, it can act as: Short‐circuit: disables all cells set in parallel
Open circuit: disables all cells set in serial
Solution: add diodes
• Serial solution with diodes
• If one cell fails and
becomes an open‐
circuit, current will
flow through the
diode.
• If one cell becomes a
short‐circuit, diode
will not affect.‐
• Parallel solution with diodes
• If one cell fails and becomes
an short‐circuit, other cells
will not be short‐circuited
because of this failed cell.
• If one cell becomes a open‐
circuit, diode will not affect.
‐
…
• This block has open‐circuit and short‐circuit protection
• Can be set in parallel with other blocks (more current)
‐ If one of the blocks have lower voltage (e.g. no light), current will
not flow through it (without damage cell risk)
• Can be set in serial with other blocks (more voltage)
‐ If one of the blocks has lower current, it will flow through diodes
(without damage cell risk)

9
• Solar Cells
– Pros
• Unlimited availability
• Low failure risk
• It can be built with common materials (Si)
Cheap
– Cons
• Low efficiency (up to 19%)
• Low power (oversized solar panels for high power consumption)
• No energy while in eclipse (shadow )
 Oversized solar panels
 Large size ⟹ Low natural resonance frequency
• Panel degradation (oversized solar panels)
• Only for inner planets missions
• Deployment mechanism complexity
• Source of single point failure
• Low efficiency solutions (1/4)

10
http://www.optoiq.com/index/photonics‐technologies‐applications/lfw‐display/lfw‐article‐display/314437/articles/laser‐
focus‐world/volume‐43/issue‐12/features/conference‐preview‐photonics‐west‐2008‐light‐it‐up‐and‐they‐will‐
comex2026‐to‐photonics‐west.html
– Multi‐junction Solar Cell
• Each layer absorbs a wavelength band
• Efficiency up to 42.3%
– Back‐surface reflectors (BSR) to reflect unabsorbed radiation and
reduce heating
• Used on Spot and Orion satellites
– Using a p+‐region (BSF) at the rear of p‐regions to enhance carrier
collector efficiency, but with high radiation fluency damage
• Used on Hubble and Envisat satellites
– Use of a textured front surface to reduce reflection from the cell
surface

11
Historic summary of champion solar‐cell efficiencies
• Fuel Cells
– Designed for manned Apollo program
– Use chemical energy converting a oxidation reaction into
electrical energy with minimal thermal changes
• Example: Hydrogen‐oxygen fuel cell, which produces
water. Useful for manned missions.
H2O fuel cell schematic
Direct Energy Conversion, Angrist, S. W., 1982

12
• Fuel cell available energy
– Available energy depends on
∆
– ∆ is “Gibbs (free) changing energy” or “free enthalpy” occurring in
the reaction,
– n is the number of electrons transferred, and
– F: Faraday constant (9.65∙104 C/mol)
• For H2O2, ∆ = ‐272,9 kJ/mol and n = 2. The resultant voltage is
1.229 V (considering an ideal cell)
• Typical hydrogen‐oxygen voltage‐current curve

13
• Fuel cell material comparison
[http://www1.eere.energy.gov/hydrogenandfuelcells/fuelcells/pdfs/fc_comparison_chart.pdf]
• Fuel cell pros
– Clean
– High efficiency
– Cheap
– Resultant substance can be useful
• In hydrogen‐oxygen cell, the output is water
• Fuel cell cons
– Low power available
– Short durability (less than a month)
• Not suitable for outer planet missions

14
• Radio‐isotope Thermoelectric Generator (RTG)
– For outer planet missions
• Cannot be used:
– Solar panels: Farther than mars, panels have to be extremely
large
– Fuel cells: Their durability is shorter than a month (in most
cases, not enough time to arrive)
• Requirements:
– High durability (Galileo took 6 years to arrive to Jupiter and
Voyager 2 took 12 years to arrive to Neptune)
• Solution: nuclear energy
• Radio‐isotope Thermoelectric Generator (RTG)
– Based on thermoelectric effect
• It is possible to generate a voltage between 2 materials if there
is a temperature gradient between them
– RTG systems use two semiconductors (p and n) with a junction to
exploit this effect
• Low efficiency (<10%) ⇒ Need to remove wasted heat
[Direct Energy Conversion, Angrist, S. W., 1982]

15
• RTG available power
– The output power depends on junction temperature
• A radioactive material heats the junction while it decays
• As it decays, the available power decreases
.
/
where τ1/2 is half‐life
– The lower τ1/2, the higher the power density
Example: Cassini‐Saturn design life was 11 years. After this time, required
power was 628 W. The only possible fuel was Plutonium.
Isotope Fuel Form Power Density (W/g) τ1/2 (years)
Polonium 210 GdPo 82 0.38
Plutonium 238 PuO2 0.41 86.4
Curium 242 Cm2O3 98 0.4
Strontium 90 SrO 0.24 28.0
• RTG pros
– Independent of orientation and shadowing
– Independent of distance from the Sun
– Low power levels for long time periods
• Example: Voyager‐2, operative from August, 1977
– They are not susceptible to radiation damage
• Example: Van Allen Belts, outer space missions…
– Suitable for missions with long eclipse time
• Example: lunar landers

16
• RTG cons
– It has to be deployed away from the satellite
• Radiation affects spacecraft electronics
• Need high temperature operation
• Example: Galileo, RTGs deployed on lengthy booms
– Careful procedures while satellite integration
– RTGs are an interference source for plasma diagnostic
equipment, which has to be deployed away.
– Space launch has a considerable
explosion risk. In this situation,
radioactive material can be dispersed
in the atmosphere.
[http://history.nasa.gov/computers/Ch6‐3.html]
• Nuclear fission systems
– Similar to ground nuclear reactors
– Started with SNAP project in 60’s: RTG developed in SNAP project
– Continued in 80’s with SP100 project
• Many reactor models: with heat pipes transporting heat to a thermo‐ionic
reactor, cooled with lithium, with liquid metal with electromagnetic pumps…
– Continued nowadays with SAFE project
• Use of Brayton Cycle gas turbine, with uranium nitrium core, up to 100 kWe
• Tested in 2001 in SAFE‐30 experiment (17.1 kW)
The Role of Nuclear Power and Nuclear Propulsion
in the Peaceful Exploration of Space, International
Atomic Energy Agency
http://web.archive.org/web/20050321055406/
http://www.spacetransportation.com/ast/presentations/7b_vandy.pdf

17
• Solar heat systems
– Original concept for the ISS
– Use of solar energy to heat a fluid and move a rotary converter
(solar dynamics) or a thermoelectric converter (solar thermoelectric)
– Conversion efficiency is 25% greater than for photovoltaic systems
– Use of a Brayton Cycle engine
– Solar light heats a fluid, helium mixed with xenon, no corrosion
– Power storage thermally using latent heat of fusion for lithium
fluoride/calcium fluoride
– Mass savings as compared
to battery technology
• Secondary power systems
– Batteries
• Provide energy while primary system is not available
• Have to be recharged while first system is available
• Example: Batteries are used while solar arrays are in shadow,
and are recharged while under sunlight

18
• Battery state of charge example
– 9 am – 9 pm orbit
[UPCSat‐1: Analysis, Design and Breadboarding of a University Picosatellite, A. Sánchez, J. Serra, A. Camps, M. Domínguez]
• Choosing the right batteries
– All batteries characteristics depend on materials
• Cell voltage
• Energy Density
• Charge efficiency
• Cycles at certain depth of discharge (DOD)
• Radiation sensitivity
• Temperature sensitivity
Material Cell
Volts
Capacity
(W∙h/kg)
Temperature
range
Cycles @ 75%
DOD
Cycles @ 25%
DOD
Ni‐Cd 1.25 30‐40 ‐20° ‐ 45° 800 21000
Ni‐H2 1.30 50‐80 0° ‐ 20° 4000 150000
Ag‐Zn 1.10 60‐70 0° ‐ 50° 100 3500
Ni‐MH 1.2 60 0° ‐ 25° 4000 130000
Li‐Ion 3.7 100‐250 ‐25° ‐ 75° 700 2700

19
• System sizing
– While operative, primary energy system should be able to
• Feed all systems
• Recharge batteries more energy than they will spend
– Batteries should be able to store more energy than which is going to
be consumed during eclipse
– Both systems should be oversized because of degradation
– We have to consider the worst situation
• Required power
– Required power at EOL (End Of Life) depends on:
• PLOADS: Power required by worst case in sunlight [W]
• Pe: Power required by worst case in shadow [W]
• te: Eclipse period [s]
• ts: Sunlight period [s]
• : Battery Charger efficiency (∼0,92)
• : Battery Discharger (regulators, buses…) efficiency (∼0,9)
• : Shunt dump efficiency

20
• Solar array sizing
– Solar arrays must be sized depending on the required power at EOL:
cos
• S = Solar intensity (∼1400 W/m2)
• cosφ= Angle between Sun line and array normal
– If we are calculating for more than one array and/or geometry
is not planar, we should replace it by “Solar Panel Orientation
Coefficient”
• = End Of Life cell conversion efficiency
• F = Sum of loss factors (∼0.75)
• PF=Packaging efficiency (∼0.85): Part of solar arrays made by solar
cells, taking into account void spaces, connections…
• ASA=Solar Array Surface
• Solar panel orientation coefficient
– Defines how many solar panels are illuminated pondering
– Illumination time (% of time under light and in shadow)
– Incidence angle
– Number of panels
– Example: Cubesat with all sides with solar arrays, rotating, with
one axis fixed towards the Earth’s magnetic field
– Orbit 6 am – 6 pm (100% sun): 1,408 x 1U
– Orbit 9 am – 9 pm (73% sun): 1,528 x 1U
– Orbit 12 am – 12 pm (64% sun): 1,473 x 1U

21
• Battery sizing
– Battery capacity should be chosen according to:
• EBAT = Battery capacity [A∙h]
• NB = Number of batteries
• NC = Number of cells
• VD = Cell discharge voltage [V]
• DOD = Depth of discharge
• Power budget
– Study of how much power is required by each subsystem as a
function of time, and how much power is available
– Need to know
• Load power consumption and on‐off time
• Systems power consumption and on‐off time
• Battery capacity
• Available power
• Light and eclipse time
• Subsystems efficiency

22
• Power budget example
– UPCSat‐1 power budget
[UPCSat‐1: Analysis, Design and Breadboarding of a University Picosatellite, A. Sánchez, J. Serra, A. Camps, M. Domínguez]
• System sizing example (1/2)
– A satellite consumes 6 W, in a 1 h of light and 15 min of shadow orbit
• Batteries have to store more than:
6 W ∙ 15 min = 1.5 W∙h
• Solar panels have to feed
6 W continuously
1,5 W∙h / 1h of sun = 1.5 W to recharge batteries
7,5 W in total
– Neither batteries and solar panel degradation, nor system efficiency
has been taken into account  margings!

23
• System sizing example (2/2)
– If discharging system efficiency is about 90%
• Total energy stored should be 1,5 W∙h / 0,9 = 1,66 W∙h
– If planned mission time is about 3 years
• 19,2 cycles/day = 21.000 cycles
• If battery type is Ni‐Cd, at the end of the mission, DOD will be 25%
• Battery should be able to store: 1,66 W∙h / 0,25 = 6,66 W∙h
• Power control and distribution
– Array regulation:
• Available and needed power vary during the mission
• At the beginning of the mission, because of oversizing, available power is
higher than required
• It is necessary to regulate power supplied by solar arrays
– Sequential switching‐shunt regulation (S3R)
– Pulse width modulation (PWM) and filters
– Maximum peak power tracking (MPPT)
– Battery control
• Battery charge regulator
• Battery management unit
• Battery discharge regulator
– Power control and distribution
– Power regulation

24
• Battery charge regulator
– Battery charging is a extremely delicate task
– Each battery type has to be charged according to certain
parameters
• Example: Li‐ion batteries have to be charged at constant
current until cell voltage reaches 4.2V. If cell voltage exceeds
that limit, it can explode.
– Depending on charging current, battery useful life will change
– NiCd battery reliability depending on overcharge factor (K)
• Battery management unit
– Monitors battery states
• Voltage, current
• Temperature, pressure
– Provides control inputs to BCR
– Interface between data‐handling subsystem and EPS (senses telemetry)
• Battery discharge regulator
– Protects battery against high discharge currents that can damage them
• High discharge currents reduce battery useful life
• Really high discharge current can make explode batteries
• Some high currents are caused by latch‐ups and SSE
– Allow batteries to full discharge in order to improve battery capacity

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