An Interplanetary CubeSat
Mission to Phobos J. Thangavelautham1, E. Asphaug1, G. Dektor2, N. Kenia2,
J. Uglietta2, S. Ichikawa2, A. Choudhari2, M. Herreras-Martinez2, S. Schwartz2
1University of Arizona
2Arizona State University
LOGIC Mission Concept to Phobos 6U CubeSat Hosted payload science mission to
Phobos 2-3 year mission life Thermal and visible camera payload Impulsive maneuver for Mars
capture 7 month science mission, with 5+
flybys Product of a 1-year graduate
capstone interdisciplinary space systems design class.
Unravelling Phobos Previous Science Missions
Viking, Phobos 1 , Phobos 2 and Fobos Grunt
Mars assets: MRO, MSL, MER, MEX, MOM Planned Mars missions:
Sample return missions : ALADDIN, MMSR, PHOOTPRINT, HALL
Imaging & spectroscopy : MPADS, OSIREX II, PCROSS, Phobos surveyor & PANDORA
High resolution imaging: PADME and MERLIN
Why Thermal Images of Phobos ? Contribute to understanding geophysical and
evolutionary properties Limited Phobos thermal and visible data
Mars Express: 3 m/pixel best Phobos 1: Failed Mission Phobos 2: 37 Images @ 40m/pixel Phobos Grunt: Lost Contact
Phobos: ISRU and Gateway to Mars
Phobos is a critical transit point to Mars
Phobos theorized to have surface contents comparable to Mars
Abundance of water ?
What are the risks ? Where are the best locations ? How do we get started ?
Primary & Secondary Mission Requirements Primary Objective: Understand the geophysical properties of
Phobos Thermal images <100m resolution 50% coverage
Composition Feature size & distribution
Secondary Objective : Understand the evolution of Phobos Visual images <10m resolution 50% coverage
Striation and crater characteristics Superposition of features
What Is Different about this Mission Concept ?
What can we do with a JPL MarCo class CubeSat but with a relaxed schedule, science focus ?
Avoid the high delta-v required for Phobos orbit. Targeted, multiple flyby’s Smarts: Visual Navigation due to limits with
Phobos ephemeris.
Aerobraking Impact
∆Vavg = -2.5 m/s • Duration: 240 Days Altitude: 120 km • Atmospheric Density: 10-8 kg/m3
• Drag Force: 0.166 N
Shallow trajectory over 6-8 months.
Aerobraking Comparison
Periapsis Altitude
Initial Apoapsis
Final Apoapsis
ΔV Duration Mass/Area
MGS 110-120 km 53 000 km 450 km 1250 m/s ~ 8 Months 83 kg/m2
Mars
Odyssey 100-110 km 30 700 km 200 km 1090 m/s 2.5 Months 34 kg/m2
MRO 98 – 195 km 48 000 km 450 km* 1190 m/s ~5 Months 52 kg/m2
LOGIC 110-120 km 226 0000 km 9520 km 690 m/s ~6-9 Months 22 kg/m2
Less ambitious than previous attempts
Trajectory Correction Maneuvers
Maneuver Time ∆V Magnitude
TCM 1 Launch + 45 d 32 m/s TCM 2 Launch + 90 d 2 m/s TCM 3 Mars - 60 d 0.5 m/s TCM 4 Mars - 14 d 0.1 m/s TCM 5 Mars – 1 d 0 -2.0 m/s
Desired Intercept • 120 km Altitude
5 major TCMs during transit.
Intercept Frequency • Time Interval: 7 Months • Encounters < 94 km: 34 Dependent on Initial Conditions: • Phobos Position at Start of Intercept
Phase • Intercept Apoapsis • Intercept Inclination
Frequency of Intercepts with Phobos
Chance encounters with Phobos. 3 Close intercepts in 7 months.
Chassis Internal Volume (cm3)
Mass (kg)
LOGIC 6U 8000 1.5
Pumpkin 6U 7700 1.1
ISIS 6U 7200 1.0
Chassis Trade Study and Down Selection
Integrated fuel tank 2400cc propellant Integrated deployment rails Integrated thruster cavity
Required Components for a Successful Mission
Top Level Requirement Component
Thermal data <100 m/pixel
FLIR Tau 2
Visible Image coverage < 10 m/pixel of ½ of Phobos
E2V Cires Camera
Alternate Requirement Component
Capture into the Mars system
MPS-130 Green Monopropellant
Thruster
Communicate ½ GB of data minimum
IRIS V2 Transponder & ISARA Reflect Array
• HaWK Solar Arrays
• Dual Gimbals • Reaction wheel solar
desaturation • Variable modes
• X-Band Reflect Array
• Hybrid on back of HaWK array
• Similar concept – ISARA (Ka-band)
Hybrid solar array and reflectarray
Attitude Determination and Control
Requirement
Pointing Accuracy +- 1 degree (Science)
Slew Rate 1.0 deg / sec (Science)
Stabilization
Shall stabilize the spacecraft for stable communication, power generation, and attitude tracking
Desaturation Shall keep the controllability through the mission
Accurate 1 degree pointing needs for science, comms and tracking.
Attitude Determination and Control Specifications XACT MAI-400 IACDS-100
Mass 0.91 kg 0.694 kg 0.35 kg
Volume 10 x 10 x 5 cm 10 x 10 x 5.59 cm 9.5 x 9.0 x 3.2 cm
Power (Stand-by) 0.03 W 0.87 W 0.15 W
Power (Nominal) 2.14 W 4.23 W 0.5 W
Power (Max. Torque) 5.55 W 8.47 W 1.8 W
Radiation Tolerance 16 krad N.S. N.S.
Max. Momentum 15 mNms 9.351 mNms 1.5 mNms
Max. Torque 4 mNm 0.635 mNm 0.087 mNm
Pointing Accuracy +-0.007 deg +- 1 degree +- 1 degree
BCT XACT overall best solution, flight heritage*, all in one solution.
XACT
Mass 0.91 kg Volume 10 x 10 x 5 cm Power (Stand-by) 0.03 W Power (Nominal) 2.14 W Power (Max) 5.55 W Radiation Tol. 16 krad Max. Momentum 15 mNms Max. Torque 4 mNm Pointing Accuracy +-0.007 deg
All in one solution, will be combined with thruster for desaturation.
Communications Subsystem
DSN compatibility EIRP - 35 dB UHF relay (investigation in progress)
Scenario (Date rate)
Distance (million km)
Downlink
Worst 300 6.4 kbps
Nominal 250 9.3 kbps
Best 200 14.5 kbps
Link budget to Earth (X band)
Iris V2 X band Transponder
NASA JPL (MarCO)
Integrated (solar array and) X-band Reflectarray with Iris v2 for Interplanetary missions
Solar Panel & Antenna Properties Specification X-band reflectarray
Mass 0.5 kg
Gain 31.69 dB (single array) 34.7 dB (double array)
1 – estimated mass of only electronics (no chassis) 2 – assuming both arrays can be synchronized
Specification MMA E-HaWK (3 panels/ wing)
Mass 0.85 kg
Power 44 W OAP
Total area of 12U x 2U shared by Reflectarray and solar panels
High gain X-band Reflectarray NASA JPL (MarCO)
Reflectarray Antenna Analysis
Efficiency Gain (dB)
55% 30.35
75% 31.69
Number of deployable Gain (dB)
1 31.69
2 34.7
Off-nominal case – Tumbling
Parameter Worst Nominal Best
Relative Power -11.7 dB -15.7 dB -25.7 dB
Tolerable theta +/- 7o +/- 8o +/- 12o
Tumbling after ejection – link can be closed irrespective of theta
MarCO reflectarray antenna gain pattern
Note – Nulls have been ignored while calculating tumbling budget
Command & Data Handing
Parameter SpaceCube MINI CHREC Space Micro
Processor Virtex 5QV – Space
Xilinx Zync – 7000 SoC
Space Qualified Yes Yes
Rad tolerance Upto 700 krad 30 krad
Frequency 450 MHz 1000 MHz
IPS 2000 MIPS 2500 MIPS
Power 5-10 W 1.54 – 2.86 W
Operating Temp -55 to 125 oC -40 to 85 oC
Space Heritage IPEX, HyspIRI11 CeREs, STP-H5/ ISEM
SpaceCube MINI : Satisfies the requirements
SpaceCube MINI
CHREC Space Micro 1 – Mission is yet to be launched
Interface Architecture
Multiple RS-422 and LVDS interfaces and Aeroflex FPGA watchdog
Subsystem Interface/ Protocol
X-band transponder
SPI
ADCS RS-422
Visual Camera RS-422
UHF Radio RS-422
Thermal Camera LVDS
Propulsion SPI
EPS SPI
Power System Critical Components Performance in Mars of E-HaWK (3 panels per wing)
Pmax at 40°C, BOL
33 W
Pmax/A at 40°C, BOL
148 W/m2
Pmax at 40°C, EOL 24 W
Primary Battery Low self-discharge rate (less than 1% a year at 20°C) Nominal voltage of 2.9V
Nominal capacity of 25 Ah (72.5Wh) Secondary Battery Self-discharge rate (2% a year)
Nominal voltage of 12V
Autonomous heater system
Power Systems Overview Main Features of Power Subsystem
PPT Architecture
MMA Solar Panels (12Ux2U total area)
Max DOD of 36% in critical scenarios
Deployment Failure Mitigation Routine allows corrections during 4h
Major Components with remarkable space heritage
Solar PV system with heritage from JPL’s ISARA mission.
Subsystem Power Usage
COMMS 0 CPU 5 PROPULSION 0 POWER 0.5
EPS 2
ADCS 2.1 PAYLOAD Thermal Camera 0 Visual Camera 0 THERMAL CONTROL
5
Total 15 Margin +38 %
Low Power - Hibernate Subsystem Power Usage
COMMS 8 CPU 5 PROPULSION 0 POWER 0.5
EPS 2
ADCS 2.1 PAYLOAD Thermal Camera 0 Visual Camera 0 THERMAL CONTROL
0
Total 18 Margin +27 %
Comms Mode - Receive Subsystem Power Usage
COMMS 24 CPU 5 PROPULSION 0 POWER 0.5
EPS 2
ADCS 2.1 PAYLOAD Thermal Camera 0 Visual Camera 0 THERMAL CONTROL
0
Total 34 Margin -40 %
Comms Mode - Transmit
Positive margin for hibernate, comms. receive. Negative margin for comms transmit to earth. Rely on battery + PV to perform comms transmit
Subsystem Power Usage
COMMS 0 CPU 5 PROPULSION 0 POWER 0.5
EPS 2
ADCS 2.1 PAYLOAD Thermal Camera 1.2 Visual Camera 1.5 THERMAL CONTROL
5
Total 17 Margin +29 %
Science Phase
Subsystem Power Usage
COMMS 0 CPU 5 PROPULSION 15 POWER 0.5
EPS 2
ADCS 2.1 PAYLOAD Thermal Camera 0 Visual Camera 0 THERMAL CONTROL
0
Total 25 Margin -5 %
Insertion Orbit
Positive margin for long science phase. Slight negative margin during insertion orbit.
Thermal Environment
Heat Load(W/m2) Q incident (W/m2) Emittance Absorptivity Q Absorbed
(W/m2) Solar Flux 640 0.55 0.35 223 Mars Albedo 192 0.55 0.35 67 Phobos Albedo 45 0.55 0.35 15 Mars IR 426 0.55 0.35 234 Total Maximum 540
Heat Load(W/m2) Q incident (W/m2) Emittance Absorptivity Q Absorbed
(W/m2) Solar Flux 490 0.55 0.35 172 Mars Albedo 147 0.55 0.35 51 Phobos Albedo 34 0.55 0.35 12 Mars IR 315 0.55 0.35 173 Total Minimum 410
( )absorbed solar Mars Phobos MarsIRQ Q Q Q Qα ε= + + +
Thermal Subystem Worst hot case Scenario •External Heat load of 541 W/m2
•Batteries dissipate 5W . •EPS and computer dissipate 5 W •Communication system dissipate 26 W and work for 4 hrs Simulation done for 2 hrs
Heat Accumulation due to Insertion Burn •Thrusters fire for 25 minutes •Maximum Nozzle Temperature of 1000 Kelvin assumed •Propellant dissipates 8 W •Simulations run for 30 minutes
Our Destiny Out There
The ability to probe and prospect for human mission stepping stones is critical.
The practicality of using Phobos as a transit base to Mars needs to be fully understood
There are important, yet high-risk, high-reward science questions that can give insight into the formation of the Martian system.
A series of CubeSat missions is probably the most cost-effective method to achieve these goals.
Conclusions
The concept shows the potential for a cost-effective, targeted method towards answering science questions.
There are important technological risks that span propulsion and system integration. The primary mission can be accomplished without Mars
insertion orbit. Majority of the proposed technologies are now standard
components for deep space missions. Falling back to a 12U or 24U can enable one or more
redundant strings, improved situation overall
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