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AIAA Exhibition Presentation_Final
The document describes a proposed concept for a constellation of 4 CubeSats in Mars orbit to augment communication relay capabilities for assets on the Martian surface and in orbit. Key aspects of the concept include using the CubeSats to relay data between landed assets and orbiters using UHF bands, and between the CubeSats and Earth using Ka-band. Mass and power budgets have been analyzed to demonstrate the feasibility of the concept. Future work areas are identified such as further analyzing data volume, thermal, trajectory, and cost aspects.
AIAA Exhibition Presentation_Final
- 1. Rohan Deshmukh Swapnil Pujari Advisor:Professor David Spencer
- 2. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work AboutUs Rohan Deshmukh Swapnil Pujari School Georgia Institute of Technology Georgia Institute of Technology Major Aerospace Engineering Aerospace Engineering Hometown Fairfax, VA Alpharetta, GA Year 3rd Year 3rd Year 5/17/2015 2
- 3. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work ConceptMotivation 5/17/2015 3 Credit: Charles Edwards, JPL
- 4. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work MarsCommunication Architecture 2025 Dedicated Communications Orbiter (1-3 satellites) 5/17/2015 4 Credit: Charles Edwards, JPL
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- 8. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work Structure/Sizing DeploymentMechanisms: 1. < 3U CubeSat: PPOD (Poly-Pico Satellite Deployer) 2. > 3U CubeSat: CSD (Canisterized Satellite Dispensers) o CSD Limitations for 6U: mass of 12 kg, size of 12 cm x 24 cm x 36 cm 35.7 cm 11.8 cm 22.2 cm Stowed CubeSat Configuration CSD Deployer Mass: 7.9 kg 5/17/2015 8
- 9. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work SolarFlux Inverse Square Law • Earth Solar Flux: 1370 W/m^2 @ 1AU • Mars Solar Flux: 589.1 W/m^2 @ 1.525 AU • Power generation is proportional to the inverse square of the distance to the sun Inverse Square Law • Drives iterative process of solar array sizing 57% Reduction in Solar Flux Density 5/17/2015 9
- 10. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work PowerSubsystem Subsystem Component Name Quantity EPS Clyde Space FLEX EPS 1 CS 30 Whr Battery 1 CS Deployable Double Sided 6U Panels 4 CS 2UFixed Solar Panels 1 Clyde Space FLEX EPS Clyde Space 30 Whr Battery CS Deployable Double Sided 6U Panels: • 21 cells per face -> 42 cells on one panel -> 84 cells one side • Solar Cell Efficiency: 28.3% • Area of one solar cell: 0.00275 m2 • Total Solar Cell Area: 0.231 m2 CS 2U Body-Mounted Solar Panels 5/17/2015 10
- 11. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work SatelliteCommunication Bands UHF X-Band Ka-Band Space Communication Usage Used between landers, rovers, and orbiters Current Standard in long range communication s (rovers <-> Earth) Developing Standard Advantages Less prone to atmospheric interference Long Range Communication Greater Data Transmission Disadvantages Smaller Distance coverage Requires Line of Sight, More power Development & Testing Stages, Large Propagation Losses Common Uses Broadcast TV, Cell Phones Radar, Deep Space Network Radar, Deep Space Network 5/17/2015 11
- 12. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work Today’sMars Relay Network Testing Ka-Band 5/17/2015 12 Credit: Charles Edwards, JPL
- 13. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work Tomorrow’sMars Relay Network MAVEN Agency: NASA Launch: Nov 18, 2013 Orbit: 150 x 6,200 km elliptical 75 deg inclination Non-sun-synchronous Deep Space Link: - Band X-band - Power Amplifier 100 W TWTA - High Gain Antenna 2 m HGA (body fixed) Proximity Link: - Transceiver Electra (single string) - Protocol CCSDS Proximity-1 - Antenna Quadrifilar Helix - Forward Link - Frequency 435-450 MHz - Data Rate 8, 32, 128 kbps - Coding (7,½) Convolutional - Return Link - Frequency 390-405 MHz - Data Rate 1, 2, 4, …, 2048 kbps - Coding (7,½) Convolutional, LDPC - Other 8-bit I/8-bit Q open loop recording Suppressed Carrier Modulation Adaptive Data Rates ExoMars/TGO Agency: ESA Launch: Jan 7-27, 2016 Orbit: 400 km circular 74 deg inclination Non-sun-synchronous Deep Space Link: - Band X-band - Power Amplifier 65 W TWTA - High Gain Antenna 2.2 m HGA Proximity Link: - Transceiver Electra (dual string) - Protocol CCSDS Proximity-1 - Antenna Quadrifilar Helix (2) - Forward Link - Frequency 435-450 MHz - Data Rate 8, 32, 128 kbps - Coding (7,½) Convolutional - Return Link - Frequency 390-405 MHz - Data Rate 1, 2, 4, …, 2048 kbps - Coding (7,½) Convolutional, LDPC - Other 8-bit I/8-bit Q open loop recording Suppressed Carrier Modulation Adaptive Data Rates 5/17/2015 13 Credit: Charles Edwards, JPL
- 14. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work Telecommunications SubsystemComponent Name Quantity Telecom IRIS Transponder 1 KaPDA 1 AX5042 UHF Transceiver 1 UHF 4 Monopole Antenna (ISIS) 1 IRIS Transponder (NASA JPL): – Currently: X-Band (Rx/Tx) – Under Development: Ka Band receiver & exciter – Future Development: UHF transponder – BPSK Modulation AX5042 UHF Transceiver: • PSK Modulation – Max 600 kbps Ka-band Parabolic Deployable Antenna (KaPDA): UHF 4-Way Monopole Antenna5/17/2015 14
- 15. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work FullEquipment List Subsystem Component Name Quantity Dimensions (l x w x h) mm ADCS BCT-XACT Module 1 100x100x50 Telecom IRIS Transponder 1 400x100x100 KaPDA 1 150x100x100 AX5042 UHF Transciever 1 5x5x1 UHF 4 Monopole Antenna (ISIS) 1 98x98x7 Propulsion Aerojet MRS-142 1 100x100x100 Structures 6U Cubesat Structure 1 100x226x340 C&DH Tyvak Computer 1 83x94x10 EPS Clyde Space FLEX EPS 1 95x90x16 CS 30 Whr Battery 1 95x90x20 CS Deployable Double Sided 6U Panels 4 200x300x1.6 CS 2UFixed Solar Panels 2 200x100x1.6 5/17/2015 15
- 16. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work MassBudget Subsystem Component Name Mass (g) Contingency Total Mass (g) ADCS BCT-XACT Module 850 10% 935 Telecom IRIS Transponder 400 10% 440 AX5042 UHF Transceiver 1 10% 1.1 UHF Monopole Antenna (ISIS) 100 10% 110 KA Antenna 1000 25% 1250 Propulsion Aerojet MRS-142 1574 10% 1731.4 Structures 6U CubeSat Structure 1100 20% 1320 C&DH Tyvak Computer 55 10% 60.5 EPS Clyde Space FLEX EPS 173 10% 190.3 CS 30 W-hr Battery 260 10% 286 CS Deployable Double Sided 6U Panels 290 30% 1508 CS 2U Fixed Solar Panels 67 10% 73.7 Total Mass 7906 Total Allowed Mass 12000 Margin 52% 5/17/2015 16
- 17. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work MarsEphemeris Data Analysis Max Phase Angle (°): 46.594 Max Earth-Mars Distance (AU): 2.64 Min Earth-Mars Distance (AU): 0.41 0.00 0.50 1.00 1.50 2.00 2.50 3.00 0 5 10 15 20 25 30 35 40 45 50 58500 59000 59500 60000 60500 61000 Mars-EarthDistance(AU) SMEPhaseAngle(deg) Modified Julian Date SME Phase Angle & Mars-Earth Distance from 2020-2025 SME phase Angle (deg) Mars-Earth Dist (AU) 5/17/2015 17
- 18. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work PowerBudget Subsystem Component Name Contingency (%) Maneuvering Power Draw (mW) Safe Mode Power Draw (mW) Operation Power Draw (mW) ADCS BCT-XACT Module 10% 1479.50 740 1480 Telecom UHF Monopole Antenna (ISIS) 15% 74.01 74 1480 AX5042 UHF Transceiver 10% 90.09 90 90 IRIS Transponder 15% 601.68 602 12034 Propulsion Aerojet MPS 10% 2200.00 0 0 C&DH Tyvak Computer 10% 550.00 550 550 EPS Clyde Space FLEX EPS 10% 165.00 165 165 CS 30 Whr Battery 10% 55.00 55 55 Total (W) 5.215 2.276 15.853 Margin 291% 797% 29% Power Generated by Solar Cells (W) Worst-Case With Shading 20.416 5/17/2015 18
- 19. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work BatteryCharge Cycle: Operation Mode 5/17/2015 19
- 20. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work LinkBudget Best Case Scenario Assumptions Units Data Rate kbps Transmission Time/Orbit sec Data Throughput/ Orbit kiloBits Megabits Gigabits 4-Sat Constellation Total (Megabits) 5250.77km (periapsis MAVEN) Cubesat -> Orbiter Orbiter -> Cubesat 1.07 2.14 1267.67 1267.67 1356.54 2713.09 1.36 2.71 0.00 0.00 5.43 10.85 Latitude 0° Cubesat -> Surface Surface -> Cubesat 199.70 203.69 349.00 349.00 69695.61 71089.52 69.70 71.09 0.07 0.07 278.78 284.36 0.41AU Cubesat -> Earth Earth -> Cubesat 87.05 26985.78 4446.13 4446.13 387039.73 1.20E+08 387.04 119982.32 0.39 119.98 1548.16 479929.26 12010.34km (apoapsis MAVEN) Cubesat -> Orbiter Orbiter -> Cubesat 0.20 0.41 1267.67 1267.67 259.28 518.56 0.26 0.52 0.00 0.00 1.04 2.07 Latitude 20° Cubesat -> Surface Surface -> Cubesat 199.70 203.69 215.00 215.00 42935.69 43794.40 42.94 43.79 0.04 0.04 171.74 175.18 2.68AU Cubesat -> Earth Earth -> Cubesat 2.04 631.34 4446.13 4446.13 9054.94 2807029.86 9.05 2807.03 0.01 2.81 36.22 11228.12 Worst Case Scenario Assumptions Units Data Rate kbps Transmission Time/Orbit sec Data Throughput/ Orbit kiloBits Megabits Gigabits 4-Sat Constellation Total (Megabits) 5/17/2015 20
- 21. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work ConceptSummary • An equatorial constellation of 4 low-cost CubeSats in Mars orbit – Augments telecom relay capability for landed assets and a current/future orbiter • This concept is targeted for the 2020/2022 timeframe, deploying as a secondary payload on Mars 2020 or a future Mars orbiter 5/17/2015 21
- 22. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work ConceptSummary • Optimized Mass/Structure to fit CSD Deployer Requirements – Sufficient margin for future iterations of spacecraft development • Performed Telecommunications Link Budget: 1) Cubesat <-> Earth: Ka-Band 2) Cubesat <-> Ground: UHF 3) Cubesat <-> Orbiter: UHF • Optimized to meet power requirements of spacecraft bus – Considered body & shadow shading effects – Analysis based on Sun-Mars-Earth Angle 5/17/2015 22
- 23. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work FutureWork • Data Volume Analysis • Thermal Analysis • Trajectory/Delta-V Analysis • Risk Management Assessment • Cost Assessment • End of Life Assessment 5/17/2015 23
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- 26. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work CostBudget (FY14) Subsystem Component Name Quantity Cost ($FY14) Contingency Total Cost $ (FY14) ADCS BCT-XACT Module 1 0 10% $0.00 Telecom IRIS Transponder 1 0 10% $0.00 KaPDA 1 0 10% $0.00 AX5042 UHF Transceiver 1 5000 15% $5,750.00 UHF Monopole Antenna (ISIS) 1 4948.43 20% $5,938.12 Propulsion Aerojet MRS-142 1 0 15% $0.00 Structures 6U Cubesat Structure 1 8082.43 10% $8,890.67 C&DH Tyvak Computer 1 30000 15% $34,500.00 EPS Clyde Space FLEX EPS 1 10550 10% $11,605.00 CS 30 Whr Battery 1 3850 10% $4,235.00 CS Deployable Double Sided 6U Panels 4 14300 40% $80,080.00 CS 2UFixed Solar Panels 1 4525 10% $4,977.50 Margin 25.00% Margin Total Cost/Cubesat $194,970 Margin Total Cost of Constellation $779,881.45 5/17/2015 26
- 27. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work SolarCell Shading Analysis SME Worst Angle (46.6°) Max Shading Effects (~6°) 5/17/2015 27
- 28. Motivation/ Background Overview Flight System Technical Resource Budgets Summary Future Work FutureMissions to Mars: 2013 - 2023 J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D J F M A M J J A S O N D Orbiters: ODY MEX MRO MAVEN ExoMars TGO Landers: Opportunity Curiosity InSight ExoMars 2016 EDM ExoMars 2018 Lander ExoMars 2018 Rover Mars 2020 Rover 202320172016201520142013 20222021202020192018 ... ... ... L M 18 Nov 22 Sep ... E E 28 Sep 19 Oct L 4 Mar - 26 Mar L 7 Jan - 27 Jan ... ... 19 Oct L 7 Jan - 27 Jan M ... E 15 Feb- 12 Feb L 26 Jul - 14 Aug E 17 Jan L 5 May - 28 May ... ... ... ... Cruise' Aerobraking' Primary'Science' Phase' Funded'Extended'' Mission'Phase' L' M E' Launch' MOI' EDL' Unfunded'Extended'' Mission'NoBonal'Plan' Legend:' 5/17/2015 28 Credit: Charles Edwards, JPL
Editor's Notes
- #4 From the picture of current & future mission through 2020, the Martian Space will need plenty communication needs with Earth and potentially other assets on/orbiting Mars Final Consideration: Cubesats have very focused missions. Not enough space to accommodate communications. Develop infrastructure using off-the shelf components to provide lower costs while increasing performance.
- #5 694634main_Pres_Mars_Comm-Nav_Evolution-Mars_Society.pdf Although a little outdated considering NASA is considering to propose a Mars 2022 Telecommunication’s Orbiter (replacement for MRO). The objective of our concept is to provide a cheap means of satifying NASA’s communication needs in the future for missions to Mars. With the development of a dedicated cubesat telecom constellation, we can reduce costs NASA might accure through the development of expensive multi-million dollar satellite missions while allowing for NASA to continue expanding their science interests on Mars. Talk about costs that cubesat is cheaper than dedicated comm sat or constellation of sats charles.d.edwards@jpl.nasa.gov
- #6 MSL -> Ground asset MRO -> Comm Orbiter 2020 Rover -> Equitorial orbit
- #7 Fix the monopole design on everywhere MRO orbit
- #8 Here is a CAD model of the overall cubesat with labels to selected visible subsystems (this is the overall deployed view). The major subsystems that this concept looks over are the telecommunications and power systems. In the future, more analysis will be done to look at all subsystems.
- #9 When selecting external components for Cubesat, structure and sizing limitation must be put into consideration. Nevertheless, these constraints acted as driving factors in various trade studies in selection components (in particular EPS and Telecom Subsystems). As you can see, the Deployable/Foldable components to fit 6U limitation set by the CubeSat deployment system (CSD). Note: These CSD limitations are current limitations as NASA is currently trying to expand the CSD system to accommodate for 6U Cubesats up to 14kg.
- #10 The o refers to earth.
- #11 Using the aforementioned analysis of Solar Flux, a simple solar cell sizing was conducted in order to determine what area of solar cells were needed in order to provide power to all critical subsystems. Through the solar cell sizing, it was determined that a total solar cell area of 0.246 m^2 is required, which corresponds to a total of 90 solar cells. Nevertheless, the two CS solar panel configuration were selected to provide a total of 92 solar cells. The deployable 6U panels will have to be custom made by CS (21 cells vs 18 cell current config). Explain how the 6U body mounted is custom made as deployable. 2 of these panels. Mainly for shading
- #12 Before diving into our telecom analysis, it is important to look into the common satellite communication band that NASA is currently/plans to use in the future on Mars. (Explain left table). The right table provides a comparison of various communication bands in terms of their frequency (Ka Band high f, UHF band lowest of the three)
- #13 The table above is representative of the current telecommunication satellite relay around Mars. As the table shows, all satellites utilize X-band as their Deep Space Link (in fact it is the standard that NASA uses on DS missions). In addition for relay, UHF transceivers are utilized (UHF is the standard that NASA uses for proximity communication.
- #14 The table above is representative of the future communication relay satellite system around Mars. As we can see from the image, these future missions will continue to utilize X-Band as their DS link. Furthermore, both satellites seem to be utilizing a standardized proximity link using a JPL-standarized Electra Transceiver.
- #15 Using the two aforementioned table for current and future satellites around Mars, it can be seen that NASA is moving towards a standardization in Proximity operations around Mars through their Electra system and continuation of utilizing X-Band as their DS link. Using these as driving factors, our proposal strives to not only meet current telecom needs, but improve on them by providing alternate cost-effective ways and test novel concepts in improving telecom performance.
- #17 Mass budget here provides a preliminary estimate of the total mass of each 6U Cubesat. The mass of each component is provided (based on technical specs sheet) and a contigency factor was added in order to account of things such as custom manufacturing costs, additional weightage due to electrical wiring, harnesses, material selection, etc. As you can see, our cubesat weights around 8 kg, which is well below the max of 12 kg set by the CSD Deployer. Duto the fact that we have not sized many of the other critical systems such as thermal and propulsion, our 52% margin allows us some flexibility in future development.
- #20 The batery DoD is for the 30 Ahr CS Battery 3.7 Ahr DoD (horiz line in purple) In general, when power generated > power draw -> charge battery power generated < power draw -> eclipse time behind Mars (approx 36%) With the given DoD info from CS specs, computed approx 402 days battery lifetime. Since we will have a lower DoD this will increase the battery lifetime
- #21 Cubesat <-> Earth: Worst Case is 2% of Best Case Total Data Throughput Cubesat <-> Surface: Worst Case is 62% of Best Case Total Data Throughput CubeSat <-> Orbiter: Worst Case is 20% of Best Case Total Data Throughput
- #22 Mass/Structure: Optimized configuration meets CSD deployer requirements Telecom: Performed Link Budget to see if DTE/DFE was possible. Ka Band + KaPDA (high gain) allow for communication UHF capability to talk with other assets (mainly Electra Radio compatible) Power: Optimized to meet power requirements of s/c bus (mostly dictated by telecom subsystem in operational mode) Considered shading effects (both body and shadow) due to Ka Antenna primarily Using ephemeris data, found max solar angle and optimized Solar Cell system based off worst case scenario MAVEN and TGO selected because they are most likely to be in service in the 2020+ timeframe Leverages the Aerocapture CubeSat concept for deployment of CubeSats into a near-equatorial orbit during approach
- #23 Mass/Structure: Optimized configuration meets CSD deployer requirements Telecom: Performed Link Budget to see if DTE/DFE was possible. Ka Band + KaPDA (high gain) allow for communication UHF capability to talk with other assets (mainly Electra Radio compatible) Power: Optimized to meet power requirements of s/c bus (mostly dictated by telecom subsystem in operational mode) Considered shading effects (both body and shadow) due to Ka Antenna primarily Using ephemeris data, found max solar angle and optimized Solar Cell system based off worst case scenario
- #27 This is a preliminary cost budget of the components selected for developing the cubesat. A few assumptions were made in computing the total cost of cubesat constellation. First off, any specific component with a cost of $0 is not really free but components which are either currently under development and/or price of component is not currently available. Nevertheless, to compensate for this, a hefty margin of 25% was added when computing the overall total costs. In addition, another assumption made is that the cost only looks at the developing the cubesat itself and does not look into other costs such as launch, DSN costs, etc. For some components such as ISIS Monopole antenna, a larger contingency was added to account for currency conversion (here from euros to USD) as well as export or tax costs.
- #28 Assumptions in analysis: Cubesat is inertial Parabolic dish on right-side of front view Orbit of cubesat lies on Martian ecliptic plane and martian orbit lies on sun ecliptic plane (i.e ignore effects of solar declination angle) Once shading area went off the surface of cubesat (Max shading), the area of shading was set to be constant (i.e increasing angle did not effect power generated)