Conceptual design and architecture of turkish communication satellite turksat 6 a final presentation

TURKSAT 6A
PRELIMINARY DESIGN
REPORT
TEAM HERMES

CONCEPT OF OPERATION
• General View of Satellite

MISSION DESIGN
General Requirements For Mission
• Satellite (SC) nominal orbital location is GEO 42°
East longitude
• SC during its mission time should have a completely
stabilized GEO phase; should be able to function up to
8° inclined orbit; should be able to de orbit after end of
mission
• By end of mission life the SC shall be able to
manever to a min of 350 km above the GEO. The fuel
budget shall include this action.
• SC shall be designed to withstand orbital conditions
at GTO and GEO

MISSION DESIGN
Orbital Calculations
First Scenario / Bi Elliptic Transfer :
The bi-elliptic transfer requires much longer transfer time
compared to the Hohmann Transfer. However, bi-elliptic is
more efficient for long distance orbit transfer.
For R < 11.94 Hohhman
For R > 15.58 Bi-elliptic
R = 6.36

MISSION DESIGN
Second Scenario / Hohhman Transfer :
ΔV for orbital insertation (includes inclination change)
Definition Velocity (m/s) TOTAL*
V1 7754.8212
ΔV1 = 1532.25V2 3074.6516
Vapogee
1602.6220
Vperigee 10194.8980

MISSION DESIGN
Station Keeping of Satellite :
•1- East-West Station Keeping (Longitude) : Mainly caused by the non-
uniform gravitational field of the Earth
ΔV = 1.58 m.sec-1/yr
•2- North-South Station Keeping (Inclination) : The influence of the
gravitational attraction of the Sun and the Moon. ,
ΔV= 45.99 m.sec-1/yr
For 22,5 years due to our design margin is 50%
ΔVtotal = 1070.6299 m/s

MISSION DESIGN
End of Life :
Communication satellite completes its mission time, it will
change its orbit and this orbit will be far from 350
kilometers.
Definition Velocity ( m/s) Separated TOTAL
ΔV2
V1 3074.6123 6.40
12.7346Vperigee 3081.0123
V2 3061.9822 6.33
Vapogee 3055.6476

MISSION DESIGN
Total Δ𝑉 Table
Altitude (Km) Inclination
(Degree)
ΔV (m/s) Mass Propellant
(Kg)
0 5.23 7754.82 -
250
(GTO Perigee)
7 2440.08 -
35786
(GTO Apogee)
0 1532.25 917.645
35786 0 12.73
De orbiting Maneuver
22.578
36136 0
Station
Keeping
0 47.58 (Per Year)*22.5 Year 593.015
Error of Total
Firing
13.44 10.394
TOTAL 2629.06 1543.633

References
• Space Mechanics – Orbital Maneuvers
• http://www.cdeagle.com/omnum/pdf/hohmann.pdf
• Spacecraft System Design by Brown
• Spacecraft Dynamics and Control by Marcel J. Sidi
• Orbital Mechanics for Engineer Students by Howard
D. Curtis

Payload specifications/Design , GS
considerations
Requirements:
TURKSAT 6A SC shall have 16 active, 4 backup Ku-band, 2 active, 1 back up X-band
transponders.
TURKSAT 6A SC Ku-Band TRS frequencies bands17.300-18.100 MHz uplink, 11.700-
12.750 MHz downlink. Ku band TRS 140W or over
X band transponder output shall be 150 W or more
TURKSAT-6A SC Ku-band coverage shall have 3 regional fixed field,
X-band has single coverage area identical to Ku-Turkey.
TURKSAT 6A shall have 1 active and 1 back up X band beacon transmitter.
Telemetry, telecommand and distance measurement functions shall be over Ku
Band for the entire lifespan of TURKSAT 6A
SC shall have a redundant Ku band telemetry and beacon transmitter. Beacon shall
transmit telemetry when needed.

PAYLOAD
•Ku-Band
• On-board switching transponder subsystem
(regenerative)
• Multi beam antenna subsystem
•X-Band
• Bend pipetransponder subsystem (transparent)
• Antenna subsystem
Coverage Areas
BLOCK DIAGRAM OF COMMUNICATION PAYLOAD
(THALES ALENIA)

UPLINK & DOWNLINK FREQUENCY RANGES and
FREQUENCY PLANS
• Ku-band Mission will be done in 745 MHz
frequency range
• X-band mission will be done in 90 MHz
frequency range
Ku-band Frequency Plan
X-band Frequency
Plan

ANTENNA SIZING AND PLACEMENT
• Ku-band Mission
• 3 antenna for 3 coverage areas
• Parabolic reflectors
• Ku-West will be mounted on Earth
deck of SC
• Ku-East & Turkey will be mounted
on East deck
• Ku-East & West antennas have
multibeam offset feeders
• X-band Mission
• 1 front-fed symmetrical antenna
• Mounted on Earth deck of SC

TRANSPONDER SUBSYSTEM
Ku-band Mission
•16 active 4 redundant TWTAs with
180 W output power
•3 active, 3 redundant Reciever,
LNA, Converter
•4 AmHeris On Board Processor
•20 EPC and L/CAMP
•IMUXes, OMUXes, Switches,
Harnesses, etc
X-band Mission
•2 active 1 redundant TWTAs with
150 W output power
•1 active, 1 redundant reciever, and
beacon transmitter
•EPC, L/CAMP, IMUXes, OMUXes,
Switches, Harnesses, etc
TTC Mission
•1 active, 1 redundant TTC Reciever
and Transmitter
MASS BUDGET OF
COMMUNICATION PAYLOAD
Sub-Parts
Total
Mass [kg]
Total mass
with
package
[kg]
Ku-Band
Repeater
218,64 280,8
X-Band
Repeater
22,45 28,0625
Ku-Band TTC 7,6 9,5
Antennas 39,7 49,625
TOTAL 288,39 367,99

POWER BUDGET OF COMMUNICATION PAYLOAD
Components Quantity
Unit Power
Consumption [W]
Total Power
Consumption [W]
Ku-Band Receiver 3 9 27
Ku-Band L/CAMP 16 3,2 51,2
Ku-Band TWTA 16 265 4240
Ku-Band LNA 3 1,8 5,4
Ku-Band OBP 4 210 840
Ku-Band Repeater 5355,6
X-Band Receiver 1 7,4 7,4
X-Band L/CAMP 2 3,2 6,4
X-Band TWTA 2 230 460
X-Band EPC 2 12 24
X-Band Beacon Trans. 1 15 15
X-Band Repeater 512,8
TTC Receiver 1 10 10
TTC Transmitter 1 20 20
Ku- Band TTC 30
TOTAL Payload Power
Cns.
5898,4

LINK BUDGET ANALYSIS
* Calculations made for 3 different cities from 3 coverage areas (Berlin, Istanbul, Abu Dhabi)
for uplink and downlink

REFERENCES:
1. Brown, Charles, Elements of Spacecraft Design
2. Elbert, B.R.; Satellite Communication Applications Handbook
3. Elbert, B.R.; Introduction to Satellite Communication
4. Calcutt, D. & Tetley, L., Satellite Communications: Principles & Applications
5. Richharia, M., Satellite Communications Systems
6. Atayero, A.A., Luka, M.K, Alatishe, A.A., Satellite Link Design: A Tutorial
7. Lee, Y.M., Eun, J.W., Lee, S.P.; Conceptual Design of the Ka-band Communication Payload for
Communication,
Oceanographic, and Meteorological Satellite (COMS)
8. TESAT Brochure: “Facts, Products and Services”
9. Thales Alenia Space Product sheet: “Ku-band space Traveling Wave Tubes”
10. Thales Alenia Space Product sheet: “X-band space Traveling Wave Tubes”
11. MELCO Product Sheet: “X-Band Beacon Transmitter”
12. TESAT Product Sheet: “Ku-band TTC Transponder”
13. http://telecom.esa.int/telecom/www/object/index.cfm?fobjectid=7923
14. http://www.satellite-calculations.com/Satellite/Downlink.htm “DVB-S & DVB-S2 Downlink
budget calculator”
15. http://www.satcom-services.com/VSAT/downlink.html “Geostationary Satellite Link Budget
Calculations Downlink”
16. http://www.satsig.net/satellite-tv-budget.htm “Satellite Link Budget for digital satellite
television MPEG distribution”
17. http://www.tubitak.gov.tr/sites/default/files/1007-udhb-2013-01_ek_1.pdf “TURKSAT 6A
UYDUSU ÖN TEKNİK İSTERLERİ”

MASS AND POWER ESTIMATES
Power Budget :
P (payload) = 5893 W
P (total) = 1,1568 * 5893 + 55,497
P (total) = 6872 W
P (subsystem) = 6872-5893= 979 W
with margin %10 : 7559,2 W
Subsystem Mass
percent
age(%)
Allocate
d
mass(kg)
Structure 21 276
Thermal 4 52,57
ACS 7 92
Power 26 341,71
Cabling 3 39,43
Propulsion 7 92
Telecom - 0
CDS 4 52,57
Payload +
Antennas(50)
28 368
1314,28
Margin 20 130
On orbit dry mass 1577,13
1577,13,1 + 292,2 + 1544 (estimated value)
Total mass = 3413,33 kg
Subsystem Percent
age
Power
(W)
Thermal
control
30 293,7
Attitude
control
28 274,1
Power 16 157
CDS 19 186
Communicati
on
0 0
Propulsion 7 69
Mechanisms 0 0

Ariane 5 ECA missions
Ariane 5 ECA is designed to deliver payloads, mainly
communications satellites, weighing up to 10 t into GTO,
including the supporting structure and adaptors. With its
increased capacity, Ariane 5 ECA can handle dual
launches of very large satellites.
Typical length: 50.5 m.
Typical liftoff mass: 780 tons
Payload capacity: 10 metric
tons to GTO; up to 20 metric
tons to LEO
Industrial prime contractor:
EADS Astrium
LAUNCH VEHICLE

The Guiana Space Center is located
in French Guiana, a French Overseas
Department. It lies on the Atlantic
coast of the Northern part of South
America, close to the equator,
between the latitudes of 2° and of 6°
North at the longitude of 50° West.
It is accessible by sea and air, served
by international companies, on regular
basis. The climate is equatorial with a
low daily temperature variation, and a
high relative
humidity. The local time is GMT – 3 h.
The Guiana Space Center
Ariane launching area
5°13’56’’ North
52°46’32’’ West
LAUNCH VEHICLE

Required Tests for
Satellite
SC Mass
(kg)
Launcher
Interface
Diameter
(mm)
1st
Fundamen
tal
Lateral
Frequency
(Hz)
Transvers
e Inertia
Wrt
separation
plane
(kg.m2)
Longitudin
al
Frequency
(Hz)
< 4500 < Ø2624 ≥ 10 ≤ 50,000 ≥ 31
Ø2624 ≥ 9
Structural Frequency
Accelerati
on (g)
Longitudinal Lateral Additional
Line Load
(N/mm)
Static Dynamic Static +
Dynamic
Lift – off -1.8 ±1.5 ±2 26
Aerodyna
mic phase
-2.7 ±0.5 ±2 23
Pressure
oscillation
s / SRB
-4.40 ±1.6 ±1 37
SRB
jettisoning
-0.7 ±3.2 ±0.9 0
Dimensioning Loads
• Structural Frequency
• Dimensioning Loads
• Spacecraft RF emission
• Sinusoidal vibration tests
• Thermal environment
• Acoustic vibration tests
• Shocks
• Safety Factors
LAUNCH VEHICLE

Spacecraft RF emission
The spacecraft must not overlap the frequency bands
of the L/V receivers 2206.5 MHz, 2227 MHz, 2254.5
MHz, 2267.5 MHz and 2284 MHz with a margin of 1
MHz. Our Transponders use Ku band.
Sine Frequency
range (Hz)
Qualificati
on
levels (0-
peak)
Protoflight
levels (0-
peak)
Acceptanc
e
levels (0-
peak)
Longitudin
al
2-5
5-50
50-100
12.4 mm
1.25 g
1 g
12.4 mm
1.25 g
1 g
9.9 mm
1 g
0.8 g
Lateral 2-5
5-25
25-100
9.9 mm
1 g
0.8 g
9.9 mm
1 g
0.8 g
8.0 mm
0.8 g
0.6 g
Sweep rate 2 oct/min 4 oct/min 4 oct/min
Sinusoidal vibration tests
LAUNCH VEHICLE

Thermal environment
LAUNCH VEHICLE
Thermal conditions under the fairing or the SYLDA 5

Acoustic vibration tests Shocks
SC tests Qualification Protoflight Acceptance
Fact
ors
Duration/
Rate
Factors Duration/
Rate
Factors Duration/
Rate
Static 1.25 N/A 1,25 N/A N/A N/A
Sine
Vibrations
1.25 2 oct/min 1,25 4 oct/min 1.0 4 oct/min
Acoustics +3
dB
or 2
120s +3 dB
or 2
60s 1.0 60s
Shock +3
dB
or
1.41
N/A +3 dB
or 1.41
N/A N/A N/A
Safety Factors /
Test Factors, Rate
and Duration
The satellite has to
provide
The values which are
in the side table.
LAUNCH VEHICLE

Tests and Companies
Tests Recommended
Company
Secondary Company Tertiary
Companies
Structural
Frequency
ESA Labs
National Space
Organization ( NSPO )
Labs
CSA Engineering
Company
Dimensioning Loads Tec-Ease Company
Spacecraft RF
emission
Raytheon Company
Sinusoidal vibration
tests
Delserro Engineering
Solutions (DES)
Thermal
environment
Axelspace Company
Acoustic vibration
tests
Data Physics
Corporation
Shocks ATA Engineering
Company
LAUNCH VEHICLE

LAUNCH VEHICLE
Adapter selection
Estimated satellite mass 3932,9 kg ; around 4000 kg
so the selection of adaptor is required strength which has to be more than
4000kg . Adapter 1194mm provides durability up to 7000 kg mass value. The
maximum mass of the adapter system is 165 kg.
(PAS 1194VS –Launch Vehicle Adapter)

LAUNCH VEHICLE
Satellite Launch Cost
Cost per kg for Arian 5 ECA is $ 15.000
The estimated kg for the satellite
3932,9 kg is around 4000 kg.
So : 4000 * 15.000 = $ 60.000.000

Transportation to The Guiana Space Center
Rochambeau international
airport
Cayenne harbor The Pariacabo docking area
Rochambeau international airport
is located near Cayenne, with a
3200 meters runway adapted to
aircraft of all classes and
particularly to the Jumbojets.
Cayenne harbor is located in
the south of the Cayenne
peninsula in Dégrad-des-
Cannes. The facilities handle
large vessels with less than 6
meters draught. The port is
linked to Kourou by 85 km
road.
The Pariacabo docking area
is located on the Kourou river,
close to Kourou city. This
facility is dedicated to the
transfer of the launcher
stages and satellites by
Arianespace ships and is
completely under CSG
responsibility. The docking
area is linked to EPCU by a 9
km road.
Prefferred Transportation way is Rochambeau international airport.
LAUNCH VEHICLE

LAUNCH VEHICLE
• http://www.tec-ease.com/about.php
• http://www.intecs.it/eng/mercati_dettagli_schede.asp?
ID_SCHEDA=25&ID_MERCATO=7
• http://www.csaengineering.com/about-us/
• http://www.raytheon.com/capabilities/products/labs/
• http://www.desolutions.com/about-us/
• http://en.axelspace.com/about/
• http://www.ata-e.com/about-ata-engineering
• http://www.nspo.org.tw/services/en/index.htm
• http://www.esa.int/esapub/bulletin/bullet92/b92eaton.htm
• http://www.dataphysics.com/applications/vibration-testing-and-shaker-
testing/satellite-testing.html
• http://www.esa.int/Our_Activities/Launchers/Europe_s_Spaceport/About_Fr
ench_Guiana
• http://www.turksat.com.tr/
References:

Budget & Equipment List Mass and PowerTotal price ($)
ADC System 3.605.000,00
OBDH System 505.000,00
Propulsion System 2.175.000,00
Payload 2.000.000,00
Structure 1.750.000,00
Electrical Power
System
7.920.000,00
Thermal Control S. 2.650.000,00
Mission Ops. 90.000,00
Management and
Engineers
6.300.000,00
Launch Vehicle 60.000.000,00
Unpredictable
expenses
4.000.000,00
Total 90.145.000,00
Total Mass
(kg)
Power
Cons. (W)
ADCS 62,12 145,9
OBC 26 60
Propulsion 85,4 62
Communic
ation 259,04 5893
Structure 310 0
Thermal
Control 57,22 200
Power 231,58 0
Total 903,22 6360,9
Estimated Values
On orbit Dry Mass : 1314,28 kg
Power Consumption: 6872 W

Tracking:
To record the data transmitted by satellite and send commands,the
receving antenna must follow and continually point directly at the
satellite in azimuth and elevation
Telemetry:
The science of gathering information at some remote location and
transmitting the data to convenient location to be examined and
recorded.
Telecommand:
The science of commanding a remote device from a convenient location
Standards:PCM TC,PKT TC,AOS TC

Osi model:
The Open Systems Interconnection model (OSI) is a conceptual
model that characterizes and standardizes the internal functions of
a communication system by partitioning it into abstraction layers. The
model is a product of the Open project at the International
Organization for Standardization (ISO).
The model groups communication functions into seven logical layers. A
layer serves the layer above it and is served by the layer below it.

Telecommand &Telemetry overview

Ku-Band TTC Transponder:
Frequency range:
Uplink: 12.500 – 14.500 MHz
Downlink: 10.700 – 12.500 MHz
We use:
Uplink: 13.00MHz Downlink: 12.10 MHz
Spacecraft Transmitter Power Output: 2W
For Downlink: 12.10 MHz 3163-05 (Ku Band)

esco Technologies company
For Downlink: 12.10 MHz 3163-05 (Ku Band)

3163-05: D=10.02 CM =0.0102 M
Gain=20.1 dBi

Downlink Telemetry Budget:
Parameter: Value: Units:
Spacecraft:
Spacecraft Transmitter Power Output: 2.0 watts
In dBW: 3.0 dBW
In dBm: 33.0 dBm
Spacecraft Transmission Line Losses: -1.0 dB
S/C Connector, Filter or In-Line Switch Losses: 0.0 dB
Spacecraft Antenna Gain: 2.0 dBiC
Spacecraft EIRP: 4.0 dBW
Downlink Path:
Spacecraft Antenna Pointing Loss: -1.0 dB
Antenna Polarization Loss: -1.5 dB
Path Loss: -154.2 dB
Atmospheric Loss: -2.2 dB
Ionospheric Loss: -0.2 dB
Rain Loss: 0.0 dB
Isotropic Signal Level at Ground Station: -155.1 dBW
Ground Station:
------- Eb/No Method -------
Ground Station Antenna Pointing Loss: -2.0 dB
Ground Station Antenna Gain: 20.1 dBiC
Ground Station Transmission Line Losses: -1 dB
Ground Station LNA Noise Temperature: 125 K
Ground Station Transmission Line Temp.: 290 K
Ground Station Sky Temperature: 450 K
G.S. Transmission Line Coefficient: 0.7943
Ground Station Effective Noise Temperature: 542 K
Ground Station Figure of Merrit (G/T): -8.2 dB/K
G.S. Signal-to-Noise Power Density (S/No): 63.3 dBHz
System Desired Data Rate: 9600 bps
In dBHz: 39.8 dBHz
Telemetry System Eb/No: 23.5 dB
Telemetry System Required Bit Error Rate: 1.00E-06
Telemetry System Required Eb/No: 18 dB

Uplink Command Budget:
Parameter: Value: Units:
Ground Station:
Transmitter Power Output: 100.0 watts
In dBW: 20.0 dBW
In dBm: 50.0 dBm
Transmission Line Losses: -3.0 dB
Connector, Filter or In-Line Switch Losses: -1.0 dB
Antenna Gain: 13.5 dBiC
Ground Station EIRP: 29.5 dBW
Uplink Path:
Ground Station Antenna Pointing Loss: -1.0 dB
Antenna Polarization Losses: -4.0 dB
Path Loss: -154.2 dB
Atmospheric Losses: -3.0 dB
Ionospheric Losses: -1.0 dB
Rain Losses: 0.0 dB
Isotropic Signal Level at Ground Station: -133.7 dBW
Spacecraft:
------- Eb/No Method -------
Spacecraft Antenna Pointing Loss: 0.0 dB
Spacecraft Antenna Gain: 59.0 dBiC
Spacecraft Transmission Line Losses: -1.0 dB
Spacecraft LNA Noise Temperature: 500 K
Spacecraft Transmission Line Temp.: 270 K
Spacecraft Sky Temperature: 290 K
S/C Transmission Line Coefficient: 0.7943
Spacecraft Effective Noise Temperature: 786 K
Spacecraft Figure of Merrit (G/T): 29.0 dB/K
S/C Signal-to-Noise Power Density (S/No): 124.0 dBHz
System Desired Data Rate: 9600 bps
In dBHz: 39.8 dBHz
Telemetry System Eb/No: 84.1 dB
Telemetry System Required Bit Error Rate: 1.00E-06
Telemetry System Required Eb/No: 17.0 dB
System Link Margin: 67.1 dB

MISSION OPERATIONS
Requirements:
• MOS shall support a mission life-time of 15 years.
• MOS will be designed to support 1.5 times the mission life-time.
• MOS shall update and maintain documents including but not limited to
the Command Dictionary, Telemetry Dictionary, and Flight Rules,
inherited from the SC development phase throughout the life of the
mission.
• MOS shall maintain the mission fault tree throughout the operational
mission.
• All data products and operations reporting shall contain Coordinated
Universal Time (UTC) time-tagging with an absolute knowledge of +/-0.6
seconds.
• All MOS processes shall include at least 20% operational margin (meaning
20% of the time allocated to do a process shall be margin).

MISSION OPERATIONS
• LEOP manoeuvres to be operated from Kourou
• Main procedures in LEOP (Launch and Early Orbit
Phase):
– Station-keeping
– In-orbit Tests (all subsystems)
• After full integration into orbit, operations to be
transferred to Gölbaşı Ground Station in Ankara
• Back-up station located at a different location in
Ankara
• Station-keeping to be carried out on every third
Monday (to meet 50% margin-of-safety).

MISSION OPERATIONS
Table: Ground Segment O&M Requirements by Functional Area
or Subsystem [Adapted from: Ebert, B. (2001). Table 11.1. The Satellite
Communication Ground Segment and Earth Station Handbook]
Equipment Characteristics Operation
Requirements
Maintenance Requirements
Antenna
system
Fixed pointing Check Operation Periodic inspection and
protective care
RF terminal
electronics
TWT High-voltage
operation,
redundancy
switching
Performance measurement and
periodic inspection; amplifier
replacement
Baseband
processing
and
multiplex
Matrix
switches
Identify failure
modes and poor
performance
Check full range of operations
yearly

MISSION OPERATIONS
Table: Preventive Maintenance Plan Structure [Adapted from:
Ebert, B. (2001). Table 11.2. The Satellite Communication Ground Segment
and Earth Station Handbook]
Major
Section
Type of Preventive
Maintenance
Frequency Typical Activity
Outdoor RF
equipment
Inspection and
performance
verification
Daily Correct beam alignment and
gain
Baseband
section
Monitoring of
performance
Continuous Measuring acquisition time,
bit error rate (BER), lost
frames and packets, time
delay, dropped calls, etc.
Power and
UPS
Monitoring of UPS,
power and HVAC
systems
Continuous Checking through centralised
monitor system and taking
action based on facility
design

MISSION OPERATIONS
Table: Corrective Maintenance Acitivites [Adapted from: Ebert, B.
(2001). Table 11.3. The Satellite Communication Ground Segment and
Earth Station Handbook]
Symptoms Possible Action Performed by Support Equipment or System
Required
Loss of RF link Restore pointing
toward satellite
RF maintenance
(electronic/mechani
cal engineering)
Antenna control unit (ACU)
adjustment; manual antenna mount
adjustment; repair or replacement
of motors, servo system or ACU
RF
interference
Check
polarisation and
adjust for
minimum
interference
Operations Spectrum analyser and power
metre
Unstable
service
performance
Troubleshoot poor
link BER or
synchronisation;
replacement of
circuit boards or
entire unit
Baseband engineer
or datacomm
specialist
BER test set, spectrum analyser,
end user equipment

MISSION OPERATIONS
References:
• BEST PRACTICES WORKING GROUP, 2003. SATELLITE MISSION
OPERATIONS BEST PRACTICES. SPACE OPERATIONS AND SUPPORT
TECHNICAL COMMITTEE AIAA. Retrieved from:
http://www.spaceops.org/images/spaceops/SOSTC-BestPractices.pdf
• Satellite Services. Türksat AŞ. Retrieved from:
http://www.turksat.com.tr/fr/satellite-services
• Chatel, F. (2011). Spacecraft Systems Engineering, 4th Edition. Peter
Fortescue (Editor), Graham Swinerd (Editor), John Stark (Editor).
• Ebert, B. (2001). The Satellite Communication Ground Segment and Earth
Station Handbook.

Starting date Ending date 1 2 3 4 5 6 7 8 9 # # # 1 2 3 4 5 6 7 8 9 # # # 1 2 3 4 5 6 7 8 9 # # # 1 2 3 4 5 6 7 8 9 # # # 1 2 3 4 5 6 7 8 9 # # #
OBDH S.
Selection ,Trade Study Jan 2016 Jan 2017
Buying and import Jan 2017 Jun 2017
Development Jun 2017 Jan 2019
Testing Jan 2019 Jun 2019
Propulsion S.
Selection ,Trade Study Jan 2016 Jul 2017
Buying and import Jul 2017 Jan 2018
Development Jan 2018 Jan 2019
Testing Jan 2019 Jun 2019
ADC S.
Selection ,Trade Study Jan 2016 Nov 2016
Buying and import Nov 2016 Jun 2017
Development Jun 2017 Jun 2018
Testing Jun 2018 Jun 2019
Launch Vehicle
Selection ,Trade Study Jan 2016 Jun 2017
Adapter selection Jan 2020 Dec 2020
Con Ops
settle in to the GS Jun 2020 Dec 2020
Thrm. Cntlr. S.
Development Jun 2017 Nov 2018
Testing Nov 2018 Jun 2019
Payload and Comm. S.
Structure
Testing & integration
first integration for testing Jun 2019 Jan 2020
Vibration Tests Jan 2020 May 2020
Structural Load Tests May 2020 Sep 2020
Year 1 Year 2 Year 3 Year 4 Year 5
Implementation Plan

Rare
(1)
Moderate
(2)
Certain
(3)
Negligible
(0)
0 0 0
Minor (1) 1 2 3
Major (2) 2 4 6
Fatal (3) 3 6 9
Risk Matrix
Risk Plan
The numbers inside the boxes refers to priority of the problem.

Risk Plan
Risk Prob. Impa
ct
Priori
ty
Actions
Stuck of solar panel mechanism 1 3 3 A reliable mechansim shouldhave been choosen. And an extra
secure mechansim shoul be considered.
Come short of budget 1 3 3 Engineering and backup component cost may be reduce.
Fail of a transponder 1 1 1 There are 4 backup transponders to overcome with this
problem.
Software Problems 2 2 4 All softwares on satellite should be reprogrammable from GS.
Unexpected decrease of propellant 2 2 4 With increasing the intervals of the control of the satellite
propellant expense may be decreased.
Get behind of schedule 2 1 2 If budget allow more engineers can be hired.
Rise of mass during the development
process that launch vehicle can't
handle with .
2 2 4 An alternative launch vehicle should be selected beginning of
the project.
Overheating of transponders 2 1 2 If a transponder overheated that much that can harm the
device. It can turn off and an backup transponder can be
started.
Low power input caused by
overheating of panels.
1 3 3 Power margin should have been choosen high enough to
hadnle eith that kind of problem.
Fail of some ADCS sensors 3 1 3 All the sensors should have enough backup to handle with a
fail of a sensor. Also horizon sensor is a kind of back up here.
Short term communication problems
that block imergency telecommands
2 3 6 Satellite software should be programmed to understand the
extreme situations and if nothing commanded from ground, it
can act to solve this problems itself.

PROPULSION SUBSYSTEM
APOGEE KICK MOTOR
• The satellite is moved from GTO to GEO by AKM.
• It must show high Isp, I and thrust and it should also has lighter mass of
propellant.
• Mp = Mf [exp(∆V/(g*Isp)) – 1] Total loaded propellant mass is ~950 kg.
• Total impulse (I) = F * t. Here, F is thrust and t is burn time.
• Northrop Grumman Space Technology Propulsion System Corporation TR-308.
Specific
Impulse(s) Thrust(N)
Propellant
Mass(kg)
Total
impulse (Ns)
Burn
time(s)
Engine
mass(kg)
TR -308 322 471,51 950,62 6870722,53 14571,743 4,76
S 400-12 318 420 980,21 6900786,85 16430,4449 3,6
S 400-15 321 425 970,32 6878151,58 16183,8861 4,3
Isp (%35) F (%25)
Mp
(%20) I (%15) Me (%5) Total
TR -308 10 10 10 10 8 9,9
S 400-12 8 8 8 8 10 8,5
S 400-15 9 9 9 9 9 9

REACTION THRUSTERS
• It’s a part of attitude control system, used for angular control and for
desaturation of wheels.
• Most important criterias for selection are appropriate thrust and low power
consumption.
• Mp = Mf [ exp(∆V/(g*Isp)) – 1]. 593 kg is needed for attitude maneuvering
propellant.
Moog ISP DST - 12 Engines
Thrust
(N)
Specific
Impulse(s)
Power
(kW) Propellant
Mass
(kg)
22N
Airbus 22 290 0.031 Bipropellant 0.65
22N IHI 22 295 0.115 Bipropellant 0.63
DST-12 22 297 0.022
Bipropellan
t 0.64
DST-11H 22 307 0.041 Bipropellant 0.77
DST-13 22 294 0.041 Bipropellant 0.68Isp (%50) Power (%40) Mass (%10) Total
22N Airbus 6 9 8 7,4
22N IHI 8 7 10 7,8
DST-12 9 10 9 9,4
DST-11H 10 8 6 8,8
DST-13 7 8 7 7,4

PROPELLANT INVENTORY
• The propellant is needed for not only orbital maneuvering also attitudecontrol.
• The Spacecraft is exposed to perturbations both N-S and W-E directions.
• MR = ẇOx / ẇFuel (Mixture Ratio =1).
• Wf = WP /1+MR and WOx = WP – Wf
Orbit
maneuver
De-orbiting
maneuver
Station-
keeping
Error of total
firing
Oxidizer Fuel
∆V (m/s) 1532,25 12,7346 1070,6299 13,445
Usable prop. (kg) 878,1292287 5,681592627 567,4791852 5,999215676
Outage prop.
(kg)(%1)
8,781292287 2,53203455 5,674791852 2,673719602
Trapped prop.
(kg)(%3)
26,34387686 1,127853085 17,02437555 1,190991081
Loaded error
(kg)(%0,5)
4,390646144 0,502271882 2,837395926 0,530394712
Total prop. (kg) 917,645044 22,57835214 593,0157485 10,39432107 766,6196 766,6196
Total loaded
prop. (kg)
1543,633466

PROPELLANT TANK BUDGET
• Total loaded propellant mass is about 1534 kg.
• ρb= (MR + 1)/[(MR/ρOx) +(1/ρf)] ( Bulk or Average density ) ρb = 1,0917 kg/lt.
• V = m /ρb the volume of propellant tank V = 1414,31 lt.
• The total volume of using tank is approximately is 1400-1500 lt.
• This value is very huge , so it is proper to divide 3 tanks each about 500 lt.
• The tanks must be made of TITANIUM and Cassini domes shape.
Propellant Symbol Density@20o
C
(kg/lt)
Fuel MMH 0,876
Oxidizer N2O4 1,447

PROPELLANT TANK TYPE
• Titanium bipropellant tank, typically used on the TV-Sat, TDF, Tele-X, Italsat, DFS
Kopernikus, Eutelsat-2, Turksat, Nahuel and Insat-2A to 2D spacecraft.

REFERENCES
• Brown, Charles, Elements of Spacecraft Design, 2002
• http://www.northropgrumman.com/Capabilities/PropulsionProductsandServices/
Documents/TR-308_DMLAE.pdf
• http://www.moog.com/literature/Space_Defense/ISP/Bipropellant_Thrusters_Rev
_0913.pdf
• http://trs-new.jpl.nasa.gov/dspace/bitstream/2014/37950/1/04-0799.pdf
• http://cs.astrium.eads.net/sp/spacecraft-propulsion/propellant-tanks/235-516-
litre-bipropellant-tank.html

ATTITUDE DETERMINATION AND
CONTROL SUBSYSTEM
Requirements
• Spacecraft should provide an orbital control of ± 0.05°
• Spacecraft should provide an attitude control of ± 0.1°
• Solar panels must be pointed at the Sun.
• Should be protected important satellite instruments.
• Antennas must be pointed at an Earth station for communication.
• For orbital maneuvers thrusters, spacecraft should be directed through desired
direction.
Stabilization Method
Three-axis stabilization method will be used for attitude control due to system
requirements. This method’s advantages are given in the following:
•High pointing accuracy
•Most adaptable to changing mission requirements
•Can accommodate large power requirements
•Unlimited payload pointing
•Can provide rapid maneuvering

CONTROL SUBSYSTEMDisturbances Torques
Solar Torque: The solar torque on the spacecraft is the sum of all of the forces on all elemental surfaces times
the radius from the centroid of the surface to the spacecraft center of mass.
Ts = PAL(1 +q) Ts = 5.13*10^-5 N-m
Magnetic Torque : Earth’s magnetic field interacts with the magnetic field created occurred by electronics of
the satellite.
Tm = Mbsinθ Tm = 3,26*10^-6 N-m
Gravity-Gradient Torque : Gravity gradient torque is a lower parts of the spacecraft are subjected to
exponentially higher gravity forces than the upper parts.
Tg= (𝟑μ/𝟑𝟑)*|𝟑𝟑−𝟑𝟑|𝟑 Tg = 2.84*10^-6 N-m
Aerodynamic Drag Torque: Aerodynamic drag Torque caused by the distance between center of pressure
and center of gravity, so it’s completely depends on the shape of the spacecraft . At GEO it’s effect almost
negligible.
Spacecraft Generated Torques: These torques are generally much smaller than the external torques. So, it is
also neglected.

CONTROL SUBSYSTEM
Attitude Determination
Sensors
Star Tracker
Protection: Star Tracker should be provided sun light therefore sun sensors should detect
Sun vector.
Heritage: More than a hundred of SED26 family Star Trackers are flying successfully since
back May 2002. It has accumulated 100% success in LEO, GEO and deep space missions.
Sodern SED 26

CONTROL SUBSYSTEM
Sun Sensor
ISS-D25
Requirements: SC should provide an orbital control of ± 0.05°
Should be protected important satellite instruments
Specifications: Field of view 50 deg
2 Axes
Position error < 0,3 deg
Accuracy < 0.04 deg
Usage: Totally 6 sun sensors will be used for ADCS and they are located
on all surfaces of satellite.
Coarse Sun Sensor
CSS-01
Requirements: Solar panels must be pointed at the Sun
Specifications: Field of View 120° full-angle circular field of view
Accuracy ±5° of 1-axis knowledge
Usage: 2 coarse sun sensors are located on solar panels so totally 4
sensors will be used for ADCS.
Heritage: It has flown successfully on multiple spacecraft, including
ALEXIS, HETE, MOST, CHIPSat, and STPSat-1. 20 years of flight heritage

CONTROL SUBSYSTEM
Earth Sensor
Sodern STD 15
Gyroscope
HRG 130P

CONTROL SUBSYSTEMKalman Filter
The Kalman filter was developed by Rudolf Kalman and is a recursive filter that estimates the
state of a dynamic system from a series of measurements which can be noisy. The Kalman
filter is also known as an estimator because it is used to estimate the current state in a
system based on the previous time step and a new set of measurements. Kalman filter will
be used to estimation determination for Turksat 6A.

CONTROL SUBSYSTEM
Attitude Control
Actuators
Rockwell Collins RSI 68
Requirements: Spacecraft should provide an attitude control
of ± 0.1°
Specifications: Operational Speed Range ± 6000 rpm
Angular Momentum at Nominal Speed: 68 Nms
Speed Limiter (eMF): < 7,300 rpm
Motor Torque at Nominal Speed: ± 170 mNm
Usage: 4 Rockwell Collins RSI 68 Momentum and Reaction
Wheels will be used for main actuators. And 4 22 N attitude
thrusters will be used for secondary. Also Thrusters will be
used for desaturation of wheels.
Heritage: High reliability through heritage and Space qualified subsystems (rotor, motor,
bearing unit and electronics)

POWER SYSTEM
* Solar Array Design
* Battery Design

Power Calculations
P(payload) = 6493.07 W
P(total) = 1.1568*6493.07+55.497
P(total) = 7566,68 W
P(subsytem) = 1073.61 W
h = orbit altitude in the same units as orbit radius
= 35786*10^3 m
R0 = radius of the central body ,usually Earth
= 6.3781366 * 10^6 m
G = Gravitational constant
= 6,6742867*10^24 Nm^2/kg^2-
Me = Mass of Earth
= 5.9721426*10^24 kg
P = orbit period (for GEO)
= 8.64*10^4 s
Tn = Maximum eclipse period for a circular orbit in the same units as orbit period
= 4.178104*10^3 s

Solar Array Design
Solar Cell Selection
Parameter Silicon
High-
efficienc
y silicon
Single-
junction
GaAS
Dual-
junction
III-V
Triple-junction
III-V
Status Obsolete SOA
Obsole
te
Nearly
obsolete
SOA
Efficiency (%) 12.7-14.8 16.6 19 22 26.8
Operating voltage (V) 0.5 0.53 0.90 2.06 2.26
Cell
weight(kg/m2)
0.13-0.50
0.13-
0.50
0.80-
1.0
0.80-1.0 0.80-1.0
Normalized efficiency
temperature
coefficient at 28˚C
-0.55%/C
-
0.35%/
C
-
0.21%/
C
-0.25%/C -0.19%/C
Cell thickness (μm) 50-200 76
140-
175
140-175 140-175
Radiation tolerance 0.66-0.77 0.79 0.75 0.80 0.84
Absorptance (ratio of
absorbed radiant flux
to the incident AM0
flux)
0.75 0.85 0.89 0.91 0.92
Compare of existing
solar cells
parameter,
according to
required
characteristics,
Triple Junciton III-V
is the best option.

Triple Junction
* A major advantage using these solar cells compared to silicon cells is that they deliver greater than
4 times higher voltage. Therefore, only one of Spectrolab’s multi-junction solar cells is required to
generate the same voltage as 5 Si solar cells connected in series
* Compared to typical silicon cells, these solar cells are over twice as efficient and thus will deliver
more than twice the power for the same area.
Characteristics
* Small and large cell sizes offered for optimum packing factor and cost competitiveness
* All sizes qualified for LEO and GEO missions
* Discrete Si bypass diode protection
* Performance for cells <32 cm2 is 28.3% efficiency (min. average @ max power, 28°C, AM0)
* Performance for cells >50 cm2 is 27.7% efficiency (min. average @ max power, 28°C, AM0)
* Available as CIC assembly (Cell-Interconnect-Coverglass with diode) for ease of integration or
delivered on completed solar panels

Solar Array Design Calculations
Psa = (((Pe * Tn ) /xn ) + ((Pd * Td ) /xd )) / Td
Psa=10987,72W
P0 =Pinc*ηc
P0 = 379.76 W/m
PEOL= P0* (ηrad* ηuv* ηs* ηcy* H1* ηm* ηt* ηcon* ηl* Lp)
PEOL = 271.28 W/m2
Asa = Psa / PEOL
Asa = 40.50 m2
packing factor for 59.65cm2
cell is 93%
Atotal = 43.55 m2
msa = 36.58 kg (84 mg/cm2
for triple junction III-V)
PlacementofSolarPanels
x*y=40.50m
2
; x=2.8m——-> y=14.46m
14.46/2=7.23mforeachside

Battery Design
Battery Type Selection
Performance of battery technologies in space use
Type
Electrode
materials
Cell
voltage
Specific
energy, W-
h/kg
Nickel-cadmium
Nickel oxide-
cadmium
1.25 24
Nickel-hydrogen Nickel-hyrogen 1.25 55
Lithium
Lithium-sulfur
dioxide
2.7 220
Silver-zinc
Silver peroxide-
zinc
1.55 175
Mercuric oxide
Mercuric oxide-
zinc
1.20 97
Fuel cells Oxygen-hydrogen 0.80
Comparing li-ion and
other batteries
*High Battery voltage
*Light weight, High
energy density
*Operational ability in
the low to high
temperature range
*Superior anti-leakage
characteristic
*High Power
Generation

Li-ion Batteries
Features/ Benefits
* Long cycle-life: 5 years ground storage and up to 20 years in orbit
* Long cycling capability: up to 100.000 cycles using adapted DOD
* High specific energy at cell level: up to 180Wh/kg
* Battery design including all safety devices and management systems
* Battery design compatible of all launcher mechanical spectrum
* Compliant with the ESA and NASA standards
* Excellent safety and reliability records
Technical specifications
* Battery voltage range from 12 to 100 V
* Battery Energy density: up to 135 Wh/kg
* Capacity range from 4.6 to 52.0 Ah
* Energy range from 16 to 180 Wh.

Battery Design Calculations
C=(Pn*Tn)/(xb-1*DOD*Vd)
C=143.5A-h
UsingSAFTVES180cells,threebatterieswouldberequiredfornominaloperationandfourwouldberequiredforbatteryoutcapability.
Ep=200*97=19400W-h
RegardingtoSAFTVES180typecellbatteries’specificenergy=165Wh/kg
19400/165=117.57kg
RegardingtoSAFTVES180typecellbatteries’endofchargevoltage=4.1V
97/4.1=23.65~24cellsperbattery

Command & Data handling Subsystem
C&DH System Requirements
Data Storage
Data
length
(bits)
Time 64
Spacecraft mode 16
Buttery bus voltages 256
Currents on buses 256
Comm system temperature 256
Eps temperature 256
Battery temperature 128
ADCS Data 20480
Thermal Control data 1024
WOD Total 22736
Packet Overhead wtih AX.25 192
Error Control (%30 of total) 6824
Total 29752
Estimated WOD Packet
Requirements
•WOD data should be send every 15 minute.
•WOD data should be collected every 30
second.
•All data should be stored on board at least 3
monthsFor 3 moths 919,3 Mbyte data collected. We
can take it about 1 Gbyte. SMU has two 16
Gbyte mass memory. It’s enough for 3 month
storage.
• Critical/irreversible events require two commands for one critical
event.
• Onboard computer memory shall be reprogrammable in flight.
• The fault protection system shall allow ground commanding of
enable/disable states.
• WOD data should be transmitted once at least in 15 minutes.
• There should be two memory storage units, one for backup.
• OBC should built in radiation hard or tolerant technology.

Command & Data Handling Subsystem
OBC Selection
RUAG SMU
• On-board satellite telecommand & telemetry functions such
as decoding, validation,
• authentication, decryption and distribution of commands
• On-board surveillance and reconfiguration functions including
the operation control modes of the unit
• Mass memory function for payload and housekeeping data
• SPARC V8 processor 65 MIPS
• RTEMS will be used as real time operation system.
• All the components build in with radiation hard technology.
• Reed Solomon will be used for downlink and uplink.
ADCS INTERFACES
• Gyro interfaces
• Reaction wheel
interfaces
• Star tracker interfaces
• Thruster control
interfaces
• Sun sensor interfaces
• Earth sensor interfaces
Typical properties
•Power consumption: <40 W average, < 60 W peak excluding
external loads
•Mass: 16 kg
•Dimensions: 420 (L) x 270 (H) x 276 (D) mm including mounting feet
• Reliability:
•>0,99 over a 3-year mission using class B components
•>0,97 over a 15-year mission using class S components
•Solar Array Motor Drive extension.
•Heaters 50W per line >500W total.
•Power distribution.

Error Handling
Command Handling

State Diagram of OBC
Orbit insertion Mode
It begins to operate right after the satellites seperates with rocket.
Communication system is on and waiting commands from gorund. Also
WOD data begins to be collected for first tests.Attitude control system
is on.Transponders are off. Thermall control system on.
Attitude Correction Mode
After the satellite reach it’s desired position at 42E longitude. This
mode will be activated. ADCS system works for the reach the desired
pointing accuracy for antennas and also solar panels.Communication
and thermal control systems are active but trasnponders are off.
Normal Mode
When the satellites reach its desired attitude. ADCS system working to
keep the desired attitude. Communication system sending WOD data
once every 15 minutes. All the subsystems of the satellite are working.
Safe Mode
If an crucial error occurs. Then the OBC switch to safe mode to handle
with the errors. Transponders, some parts of ADCS system are off. In
this mode ADCS systems still control satellite but with less accuracy
only allow to pointing the antennas through earth. Thermal control
system still working but margins decreased. After the error situation
solved, OBC mode switches to the Attitude Correction Mode.
Note all the errors switch OBC to safe mode. If error is minor and it can
be overcome easly. OBC tries the solve the problem in normal mode.

STRUCTURE
Requirements
Structure of spacecraft shall be designed considering 1.5 times life time of the mission.
Safety factor not must be less than 0.8
Primary Structure
7075-T651 Central Aluminium Tube ( Radius~1m)

STRUCTURE
Secondary Structure
In plate and shell configurations, a skin-stringer or honeycomb sandwich
configuration is able to transmit loads while minimizing weight. The honeycomb
section is slightly more efficient than skin stringer and less prone to buckling.
Honeycomb sections are made of two thin (0.5-1.0
mm) aluminum or composite sheets bonded on either side of a honeycomb core.

STRUCTURE
Support Structure for the external equipment
Advanced carbon fiber reinforced (CFRP) composites have been proposed as excellent
reinforcement materials for the fatigue strengthening of steel structures
CFRP – UHM Properties
It is an ultra high modulus (UHM) plate with a nominal elastic modulus of 460 GPa
and a nominal tensile strength of 1500 MPa.High stiffness, low CTE, very high cost.
We choice support structure materials CFRP – UHM, because antennas, star trackers
and reflectors so sensitive.

THERMAL CONTROL SUBSYSTEM
Need of Active Control
Only heaters were selected as an active
thermal control method. Heaters are needed
for batteries, thruster, optical sensors,
electronic components, propellant tanks.
REQUIREMENTS
EQUIPMENT
Operating Temperatures ( °
C )
Electronic equipments -20 to 70
Lİ-on batteries -20 to 60
Solar arrays -100 to 125
Reaction wheels -20 to 50
ISS-D25 sun sensor -40 to 85
Sodern SED-26 star
tracker
-30 to 60
Sodern SDT-15 horizon
sensor
-25 to 55
Antennas -40 to 85
Power supply -25 to 85
Propellant tank 10 to 50
Thruster 10 to 120
Worst-Case Orbit Conditions In
GEO orbit,
1.Worst case eclipse conditions 1.2
hours max during 24 hr orbit.
2.Two 45 day periods for GEO orbits
center around the equinoxes.
3.Array output power degrades by
several percent for the first year and
0.5% per year
4.15 to 20 hours for battery recharge
after an equinox.

THERMAL CONTROL SUBSYSTEMThermal Control Method Selections
Black and White Paints
• using absorption and emission provides the flow of heat.
Multilayer Insulation (MLI)
• most effective kind of thermal insulation.
• reduction of heat transfer due to radiation.
Heat pipes
• moving heat as a result of the laws of physics.
Heaters
• find application where internal power dissipation varies rather widely.
• are switched if the temperature falls close to the non-operation temperature limits of
the equipment.
Optical Solar Reflector
• reflects a large percentage of the incident radiation
Thermal radiators
• preferred to a mechanical cooler for reliability, life duration, simplicity and absence
of induced vibrations reasons.
Louvers
• using movable or rotating shutters over a radiating surface, have gained a wide
acceptance as highly efficient devices for controlling the temperature of a spacecraft.

Space Heritage
1.Qioptiq Space Technology OSR
35 years of Space Heritage
Space Qualified
Supply 80% of the world’s OSR’s
Bespoke solutions to meet every mission need
2.Aluminized Kapton Film
Kapton is used in applications such as the solar array and for thermal management in the United State’s
space program.
3.Kapton Flexible Polyimide Transparent Heaters
Kapton is used in applications such as the solar array and for thermal management in the United State’s
space program. They extremely versatile and operate efficiently in many environments.
4.Lord Aeroglaze Black and White Paint
Aeroglaze coatings are designed to meet specific technical demands for aviation and space applications.
Requirements such as erosion resistance, outgassing, thermal absorption and static discharge can all be
met using Aeroglaze technology. Approvals include NASA, UK MoD and Mil-std.

5.ACT Heat Pipes
ACT fabricates Constant Conductance Heat Pipes (CCHPs) to exact aerospace requirements. These
devices are manufactured under ACT’s AS 9100-2009 and ISO 9001:2008 certified Quality System.
6.Thermacore Inc.
For critical thermal management needs, including heat dissipation requirements of ≥ 40°C reduction,
communications satellite and spacecraft engineers turn to k-Core passive thermal radiators. They offer a
versatile, specialty engineered material, annealed pyrolytic graphite (APG) for reliable performance.
7.Louvers
Thermal louvers have gained a wide acceptance in the Aerospace industry as highly-efficient devices for
controlling the temperature of a satellite. Orbital's first louvers were flown in 1965. Since then, more
than 500 Orbital louver units have flown on numerous satellites, including NIMBUS-4, 5, 6 & 7; Landsat-
2, 3, 4 & 5; OAO A2 & A4; ATS-6, Viking-1 & 2; Voyager-1 & 2;NAVSTAR/GPS series; Solar Maximum
Mission; AMPTE, SPARTAN, Space Telescope, Magellan, GRO, UARS, EUVE, TOPEX, GOES, MGS, MSP,
MTSAT and TRMM.

Mass and Power Budgets (Calculated)
1.Qioptiq OSR
Mass= 0.035 kg, Power=0
2.Aluminized Kapton Film (12 layer)
Mass = 1.3 kg, Power=0
3.Kapton Flexible Polyimide Transparent Heaters
Mass= 0.0125 kg, Power=200 W
4.Lord Aeroglaze Black and White Paint
Mass= 29 kg, Power=0
5.ACT Heat Pipes
6.K-core Radiator
7. Orbital 41901 Louvers
Mass= 2.22 kg, Power=0 Total Power= 200 W
Total Mass= 55 kg
( They are acceptable for this mission.)

THERMAL SCHEME

References
• Elements of Spacecraft Design, Charles D. Brown
• http://ocw.mit.edu/courses/aeronautics-and-astronautics/16-851-satellite-engineering-fall-
2003/lecture-notes/l23thermalcontro.pdf
• http://www.google.com.tr/url?
sa=t&rct=j&q=&esrc=s&source=web&cd=1&ved=0CC0QFjAA&url=http%3A%2F
%2Fwww.dtic.mil%2Fcgi-bin%2FGetTRDoc%3FAD
%3DADA294249&ei=hsdaU_DJOKTH7AbUmICgDQ&usg=AFQjCNEyH-j-
5dhToPJW6Dy3NhOpjOxkmA&sig2=b_kBTdO-uzgium9xFocZQA
• http://www.pexa.co.uk/en/products/paints_coatings/space_optical_coatings
• http://pcos.gsfc.nasa.gov/studies/
• http://www2.dupont.com/Kapton/en_US/assets/downloads/pdf/summaryofprop.pdf
• http://www.orbital.com/SatelliteSpaceSystems/SpaceTechnicalSystems/ManufacturingServic
es/Publications/Thermal_Louvers_Brochure.pdf
• http://www.1-act.com/products/constant-conductance-heat-pipes/
• http://www.qioptiq.com/company-brochures.html
• http://www.thermacore.com/products/kcore.aspx
• http://www.nphheaters.com/products/kapton_flexible_heaters.htm
http://www.hexag.org/news/30/MASOUD_AMELI.pdf
• http://enu.kz/repository/2009/AIAA-2009-6111.pdf

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