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

TURKSAT 6A PRELIMINARY DESIGN

REPORT

TEAM HERMES

ORGANIZATION CHART

CONCEPT OF OPERATION

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 DESIGNOrbital 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

Vapogee1602.6220

Vperigee 10194.8980

MISSION DESIGN

MISSION DESIGNStation 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

http://www.cdeagle.com/omnum/pdf/hohmann.pdf

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 overX band transponder output shall be 150 W or moreTURKSAT-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 AreasBLOCK 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 PlanX-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 QuantityUnit Power

Consumption [W]Total Power

Consumption [W]Ku-Band Receiver 3 9 27Ku-Band L/CAMP 16 3,2 51,2Ku-Band TWTA 16 265 4240Ku-Band LNA 3 1,8 5,4Ku-Band OBP 4 210 840

Ku-Band Repeater 5355,6X-Band Receiver 1 7,4 7,4X-Band L/CAMP 2 3,2 6,4X-Band TWTA 2 230 460X-Band EPC 2 12 24

X-Band Beacon Trans. 1 15 15

X-Band Repeater 512,8TTC Receiver 1 10 10

TTC Transmitter 1 20 20Ku- 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 Design2. Elbert, B.R.; Satellite Communication Applications Handbook3. Elbert, B.R.; Introduction to Satellite Communication4. Calcutt, D. & Tetley, L., Satellite Communications: Principles & Applications5. Richharia, M., Satellite Communications Systems6. Atayero, A.A., Luka, M.K, Alatishe, A.A., Satellite Link Design: A Tutorial7. 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=792314. 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 WP (total) = 1,1568 * 5893 + 55,497P (total) = 6872 WP (subsystem) = 6872-5893= 979 Wwith margin %10 : 7559,2 W

Subsystem Mass percentage(%)

Allocated mass(kg)

Structure 21 276Thermal 4 52,57

ACS 7 92Power 26 341,71Cabling 3 39,43

Propulsion 7 92Telecom - 0

CDS 4 52,57Payload +

Antennas(50)28 368

1314,28Margin 20 130

On orbit dry mass 1577,13

1577,13,1 + 292,2 + 1544 (estimated value)

Total mass = 3413,33 kg

Subsystem Percentage

Power (W)

Thermal control

30 293,7

Attitude control

28 274,1

Power 16 157

CDS 19 186

Communication

0 0

Propulsion 7 69

Mechanisms 0 0

LAUNCH VEHICLE

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 tonsPayload capacity: 10 metric tons to GTO; up to 20 metric tons to LEOIndustrial 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 relativehumidity. The local time is GMT – 3 h.

The Guiana Space Center

Ariane launching area

5°13’56’’ North52°46’32’’ West

LAUNCH VEHICLE

Required Tests for Satellite

SC Mass (kg)

Launcher InterfaceDiameter

(mm)

1st Fundamen

talLateral

Frequency(Hz)

Transverse Inertia

Wrt separation

plane (kg.m2)

Longitudinal

Frequency(Hz)

< 4500 < Ø2624 ≥ 10 ≤ 50,000 ≥ 31Ø2624 ≥ 9

Structural Frequency

Acceleration (g)

Longitudinal Lateral Additional Line Load

(N/mm)

Static Dynamic Static + Dynamic

Lift – off -1.8 ±1.5 ±2 26Aerodynamic 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

atilay, 05/14/2014

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)

Qualification

levels (0-peak)

Protoflight levels (0-

peak)

Acceptance

levels (0-peak)

Longitudinal

2-55-50

50-100

12.4 mm1.25 g

1 g

12.4 mm1.25 g

1 g

9.9 mm1 g

0.8 g

Lateral 2-55-25

25-100

9.9 mm1 g

0.8 g

9.9 mm1 g

0.8 g

8.0 mm0.8 g0.6 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

Factors

Duration/Rate

Factors Duration/Rate

Factors Duration/Rate

Static 1.25 N/A 1,25 N/A N/A N/ASine

Vibrations1.25 2 oct/min 1,25 4 oct/min 1.0 4 oct/min

Acoustics +3 dBor 2

120s +3 dBor 2

60s 1.0 60s

Shock +3 dBor

1.41

N/A +3 dBor 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 CompaniesTests Recommended

CompanySecondary Company Tertiary

CompaniesStructural Frequency

ESA LabsNational 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 VEHICLEAdapter 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 VEHICLESatellite 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:

Block diagram of satellite

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 62,12 145,9OBC 26 60Propulsion 85,4 62Communication 259,04 5893Structure 310 0Thermal Control 57,22 200Power 231,58 0Total 903,22 6360,9

Estimated ValuesOn orbit Dry Mass : 1314,28 kgPower 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.

Osi layer models

Telecommand &Telemetry overview

Ku-Band TTC Transponder:Frequency range:

Uplink: 12.500 – 14.500 MHzDownlink: 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 wattsIn dBW: 3.0 dBWIn dBm: 33.0 dBmSpacecraft Transmission Line Losses: -1.0 dBS/C Connector, Filter or In-Line Switch Losses: 0.0 dBSpacecraft Antenna Gain: 2.0 dBiCSpacecraft EIRP: 4.0 dBWDownlink Path: Spacecraft Antenna Pointing Loss: -1.0 dBAntenna Polarization Loss: -1.5 dBPath Loss: -154.2 dBAtmospheric Loss: -2.2 dBIonospheric Loss: -0.2 dBRain Loss: 0.0 dBIsotropic Signal Level at Ground Station: -155.1 dBWGround Station:

------- Eb/No Method ------- Ground Station Antenna Pointing Loss: -2.0 dBGround Station Antenna Gain: 20.1 dBiCGround Station Transmission Line Losses: -1 dBGround Station LNA Noise Temperature: 125 KGround Station Transmission Line Temp.: 290 KGround Station Sky Temperature: 450 KG.S. Transmission Line Coefficient: 0.7943Ground Station Effective Noise Temperature: 542 KGround Station Figure of Merrit (G/T): -8.2 dB/KG.S. Signal-to-Noise Power Density (S/No): 63.3 dBHzSystem Desired Data Rate: 9600 bpsIn dBHz: 39.8 dBHzTelemetry System Eb/No: 23.5 dB

Telemetry System Required Bit Error Rate: 1.00E-06

Telemetry System Required Eb/No: 18 dB

System Link Margin: 5.5 dB

Uplink Command Budget: Parameter: Value: Units:Ground Station: Transmitter Power Output: 100.0 wattsIn dBW: 20.0 dBWIn dBm: 50.0 dBmTransmission Line Losses: -3.0 dBConnector, Filter or In-Line Switch Losses: -1.0 dBAntenna Gain: 13.5 dBiCGround Station EIRP: 29.5 dBWUplink Path: Ground Station Antenna Pointing Loss: -1.0 dBAntenna Polarization Losses: -4.0 dBPath Loss: -154.2 dBAtmospheric Losses: -3.0 dBIonospheric Losses: -1.0 dBRain Losses: 0.0 dBIsotropic Signal Level at Ground Station: -133.7 dBWSpacecraft:

------- Eb/No Method ------- Spacecraft Antenna Pointing Loss: 0.0 dBSpacecraft Antenna Gain: 59.0 dBiCSpacecraft Transmission Line Losses: -1.0 dBSpacecraft LNA Noise Temperature: 500 KSpacecraft Transmission Line Temp.: 270 KSpacecraft Sky Temperature: 290 KS/C Transmission Line Coefficient: 0.7943Spacecraft Effective Noise Temperature: 786 KSpacecraft Figure of Merrit (G/T): 29.0 dB/KS/C Signal-to-Noise Power Density (S/No): 124.0 dBHzSystem Desired Data Rate: 9600 bpsIn dBHz: 39.8 dBHzTelemetry System Eb/No: 84.1 dB

Telemetry System Required Bit Error Rate: 1.00E-06 Telemetry System Required Eb/No: 17.0 dBSystem Link Margin: 67.1 dB

MISSION OPERATIONSRequirements:• 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 OPERATIONSTable: 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 OPERATIONSTable: 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 OPERATIONSTable: 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/mechanical 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 2017Buying and import Jan 2017 Jun 2017Development Jun 2017 Jan 2019Testing Jan 2019 Jun 2019Propulsion S.Selection ,Trade Study Jan 2016 Jul 2017Buying and import Jul 2017 Jan 2018Development Jan 2018 Jan 2019Testing Jan 2019 Jun 2019ADC S.Selection ,Trade Study Jan 2016 Nov 2016Buying and import Nov 2016 Jun 2017Development Jun 2017 Jun 2018Testing Jun 2018 Jun 2019Launch VehicleSelection ,Trade Study Jan 2016 Jun 2017Adapter selection Jan 2020 Dec 2020Con Opssettle in to the GS Jun 2020 Dec 2020Thrm. Cntlr. S.Selection ,Trade Study Jan 2016 Jan 2017Buying and import Jan 2017 Jun 2017Development Jun 2017 Nov 2018Testing Nov 2018 Jun 2019Payload and Comm. S.Selection ,Trade Study Jan 2016 Jan 2017Buying and import Jan 2017 Jun 2017Development Jun 2017 Jun 2018Testing Jun 2018 Jun 2019StructureSelection ,Trade Study Jan 2016 Jan 2017Buying and import Jan 2017 Jun 2017Development Jun 2017 Jun 2018Testing Jun 2018 Jun 2019Testing & integrationfirst integration for testing Jun 2019 Jan 2020Vibration Tests Jan 2020 May 2020Structural Load Tests May 2020 Sep 2020integration of flight Model Sep 2020 Dec 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 PlanRisk Prob. Impa

ctPriori

tyActions

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 4,76S 400-12 318 420 980,21 6900786,85 16430,4449 3,6S 400-15 321 425 970,32 6878151,58 16183,8861 4,3

Isp (%35) F (%25)Mp

(%20) I (%15) Me (%5) TotalTR -308 10 10 10 10 8 9,9S 400-12 8 8 8 8 10 8,5S 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 EnginesThrust (N)

Specific Impulse(s)

Power (kW) Propellant

Mass (kg)

22N Airbus 22 290 0.031 Bipropellant 0.6522N IHI 22 295 0.115 Bipropellant 0.63

DST-12 22 297 0.022Bipropellan

t 0.64DST-11H 22 307 0.041 Bipropellant 0.77DST-13 22 294 0.041 Bipropellant 0.68Isp (%50) Power (%40) Mass (%10) Total22N Airbus 6 9 8 7,4

22N IHI 8 7 10 7,8DST-12 9 10 9 9,4

DST-11H 10 8 6 8,8DST-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 = ẇ Ox / ẇ Fuel (Mixture Ratio =1).

• Wf = WP /1+MR and WOx = WP – Wf Orbit

maneuverDe-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@20oC (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 MethodThree-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


Disturbances TorquesSolar 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.


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


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 degUsage: 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 SunSpecifications: 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


Earth Sensor

Sodern STD 15

Gyroscope

HRG 130P


Kalman 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.


Attitude Control

Actuators

Rockwell Collins RSI 68

Requirements: Spacecraft should provide an attitude control of ± 0.1°Specifications: Operational Speed Range ± 6000 rpmAngular Momentum at Nominal Speed: 68 NmsSpeed Limiter (eMF): < 7,300 rpmMotor Torque at Nominal Speed: ± 170 mNmUsage: 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 CalculationsP(payload) = 6493.07 W

P(total) = 1.1568*6493.07+55.497P(total) = 7566,68 WP(subsytem) = 1073.61 Wh = 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 mG = Gravitational constant = 6,6742867*10^24 Nm^2/kg^2-

Me = Mass of Earth = 5.9721426*10^24 kgP = orbit period (for GEO) = 8.64*10^4 sTn = Maximum eclipse period for a circular orbit in the same units as orbit period = 4.178104*10^3 s

Solar Array DesignSolar Cell Selection

Parameter SiliconHigh-

efficiency silicon

Single-junction

GaAS

Dual-junction

III-V

Triple-junction III-V

Status Obsolete SOAObsole

teNearly

obsoleteSOA

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 76140-175

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 CalculationsPsa = (((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)

Placement of Solar Panels

x*y = 40.50 m2 ; x= 2.8m ——-> y= 14.46m

14.46/2= 7.23m for each side

Battery DesignBattery Type Selection

Performance of battery technologies in space use

TypeElectrode materials

Cell voltage

Specific energy, W-

h/kg

Nickel-cadmiumNickel oxide-

cadmium1.25 24

Nickel-hydrogen Nickel-hyrogen 1.25 55

LithiumLithium-sulfur

dioxide2.7 220

Silver-zincSilver peroxide-

zinc 1.55 175

Mercuric oxideMercuric oxide-

zinc1.20 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 BatteriesFeatures/ 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 = 143.5 A-h

Using SAFT VES180 cells, three batteries would be required for nominal operation and four would be required for battery out capability.

Ep = 200*97 = 19400 W-h

Regarding to SAFT VES180 type cell batteries’ specific energy = 165 Wh/kg

19400/165 = 117.57 kg

Regarding to SAFT VES180 type cell batteries’ end of charge voltage = 4.1 V

97/4.1 = 23.65 ~ 24 cells per battery

1.1 kg for per cell

24*1.1= 26.4 kg for per battery

for four batteries :

26.4*4 = 105.6 kg

packing factor 60% ——> 176 kg

Command & Data handling SubsystemC&DH System Requirements

Data Storage

Data length (bits)

Time 64Spacecraft mode 16Buttery bus voltages 256 Currents on buses 256 Comm system temperature 256

Eps temperature 256Battery temperature 128ADCS Data 20480Thermal Control data 1024WOD Total 22736Packet Overhead wtih AX.25 192Error Control (%30 of total) 6824

Total 29752

Estimated WOD PacketRequirements•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 SubsystemOBC 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

Command & Data Handling Subsystem

State Diagram of OBCOrbit 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.

Command & Data Handling Subsystem

STRUCTURERequirementsStructure of spacecraft shall be designed considering 1.5 times life time of the mission.Safety factor not must be less than 0.8

Primary Structure7075-T651 Central Aluminium Tube ( Radius~1m)

STRUCTURESecondary StructureIn 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.

STRUCTURESupport Structure for the external equipmentAdvanced carbon fiber reinforced (CFRP) composites have been proposed as excellent reinforcement materials for the fatigue strengthening of steel structures

CFRP – UHM PropertiesIt 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.

STRUCTURE

Satellite Cad

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

EQUIPMENTOperating 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 year4.15 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.

THERMAL CONTROL SUBSYSTEMSpace Heritage

1.Qioptiq Space Technology OSR 35 years of Space HeritageSpace QualifiedSupply 80% of the world’s OSR’sBespoke 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.

THERMAL CONTROL SUBSYSTEM

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.

THERMAL CONTROL SUBSYSTEMMass and Power Budgets (Calculated)

1.Qioptiq OSRMass= 0.035 kg, Power=0

2.Aluminized Kapton Film (12 layer)Mass = 1.3 kg, Power=0

3.Kapton Flexible Polyimide Transparent HeatersMass= 0.0125 kg, Power=200 W

4.Lord Aeroglaze Black and White Paint Mass= 29 kg, Power=0

5.ACT Heat PipesMass= 3.96 kg, Power=0

6.K-core RadiatorMass= 18.6 kg, Power=0

7. Orbital 41901 LouversMass= 2.22 kg, Power=0 Total Power= 200 W Total Mass= 55 kg ( They are acceptable for this mission.)

THERMAL CONTROL SUBSYSTEM 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

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

Engineering

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