An Interplanetary CubeSat Mission to 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 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, , and Fobos Grunt  Mars assets: MRO, MSL, MER, MEX, MOM  Planned Mars missions:  Sample return missions : ALADDIN, MMSR, , 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  : 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.

LOGIC Mission Concept Trajectory

7 Month Mars Transfer Highly Elliptical Capture Orbit Orbit

→ Trajectory 6-8 Months Aerobraking Phobos Intercept

→ 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 Initial Final ΔV Duration Mass/Area Altitude Apoapsis Apoapsis MGS 110-120 km 53 000 km 450 km 1250 m/s ~ 8 Months 83 kg/m2 Mars 100-110 km 30 700 km 200 km 1090 m/s 2.5 Months 34 kg/m2 Odyssey 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

Desired Intercept • 120 km Altitude

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

5 major TCMs during transit. Frequency of Intercepts with Phobos

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

Chance encounters with Phobos. 3 Close intercepts in 7 months. Spacecraft

Chassis Trade Study and Down Selection

Chassis Internal Mass Volume (kg) (cm3) LOGIC 6U 8000 1.5

Pumpkin 6U 7700 1.1

ISIS 6U 7200 1.0

 Integrated fuel tank 2400cc propellant  Integrated deployment rails  Integrated thruster cavity Required Components for a Successful Mission

Top Level Requirement Component Thermal data <100 FLIR Tau 2 m/pixel

Visible Image coverage < E2V Cires Camera 10 m/pixel of ½ of Phobos

Alternate Requirement Component Capture into the Mars MPS-130 Green system Monopropellant Thruster Communicate ½ GB of IRIS V2 Transponder data minimum & ISARA Reflect Array Hybrid solar array and reflectarray

• 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) Attitude Determination and Control Requirement

Pointing Accuracy +- 1 degree (Science)

Slew Rate 1.0 deg / sec (Science)

Shall stabilize the spacecraft for stable communication, Stabilization power generation, and attitude tracking

Shall keep the controllability Desaturation 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. 5.55 W 8.47 W 1.8 W Torque) Radiation 16 krad N.S. N.S. Tolerance 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

Iris V2 X band Transponder  DSN compatibility  EIRP - 35 dB  UHF relay (investigation in progress)

Link budget to (X band) Scenario Distance Downlink (Date rate) (million km) Worst 300 6.4 kbps

Nominal 250 9.3 kbps

Best 200 14.5 kbps NASA JPL (MarCO)

 Integrated (solar array and) X-band Reflectarray with Iris v2 for Interplanetary missions Solar Panel & Antenna Properties

Specification X-band reflectarray Specification MMA E-HaWK (3 panels/ wing) Mass 0.5 kg Mass 0.85 kg Gain 31.69 dB (single array) 34.7 dB (double array) Power 44 W OAP

1 – estimated mass of only electronics (no chassis) 2 – assuming both arrays can be synchronized

High gain X-band Reflectarray NASA JPL (MarCO)

 Total area of 12U x 2U shared by Reflectarray and solar panels Reflectarray Antenna Analysis

Efficiency Gain (dB) Number of deployable Gain (dB) 55% 30.35 1 31.69 75% 31.69 2 34.7 Off-nominal case – Tumbling

Parameter Worst Nominal Best

Relative -11.7 dB -15.7 dB -25.7 dB Power

Tolerable +/- 7o +/- 8o +/- 12o theta

Note – Nulls have been ignored while calculating tumbling budget MarCO reflectarray antenna gain pattern

Tumbling after ejection – link can be closed irrespective of theta Command & Data Handing CHREC Space Parameter SpaceCube MINI Micro Virtex 5QV Xilinx Zync Processor – Space – 7000 SoC Space Qualified Yes Yes Rad tolerance Upto 700 krad 30 krad SpaceCube MINI 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 CeREs, Space Heritage IPEX, HyspIRI11 STP-H5/ ISEM 1 – Mission is yet to be launched CHREC Space Micro SpaceCube MINI : Satisfies the requirements Interface Architecture Interface/ Subsystem Protocol X-band SPI transponder ADCS RS-422

Visual Camera RS-422

UHF Radio RS-422

Thermal Camera LVDS

Propulsion SPI

EPS SPI

Multiple RS-422 and LVDS interfaces and Aeroflex FPGA watchdog Interface Architecture

*

http://deepspace.jpl.nasa.gov/galleries/goldstone/ Power System Critical Components Performance in Mars of E-HaWK (3 panels per wing) Pmax at 40°C, 33 W BOL Pmax/A at 40°C, 148 W/m2 BOL Pmax at 40°C, 24 W EOL 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.

Low Power - Hibernate Comms Mode - Receive Comms Mode - Transmit Subsystem Power Usage Subsystem Power Usage Subsystem Power Usage COMMS 0 COMMS 8 COMMS 24 CPU 5 CPU 5 CPU 5 PROPULSION 0 PROPULSION 0 PROPULSION 0 POWER 0.5 POWER 0.5 POWER 0.5 EPS 2 EPS 2 EPS 2 ADCS 2.1 ADCS 2.1 ADCS 2.1 PAYLOAD PAYLOAD PAYLOAD Thermal Camera 0 Thermal Camera 0 Thermal Camera 0 Visual Camera 0 Visual Camera 0 Visual Camera 0 THERMAL THERMAL THERMAL 5 0 0 CONTROL CONTROL CONTROL Total 15 Total 18 Total 34 Margin +38 % Margin +27 % Margin -40 %

 Positive margin for hibernate, comms. receive. Negative margin for comms transmit to earth. Rely on battery + PV to perform comms transmit Insertion Orbit Science Phase

Subsystem Power Usage Subsystem Power Usage COMMS 0 COMMS 0 CPU 5 CPU 5 PROPULSION 15 PROPULSION 0 POWER 0.5 POWER 0.5 EPS 2 EPS 2 ADCS 2.1 ADCS 2.1 PAYLOAD PAYLOAD Thermal Camera 0 Thermal Camera 1.2 Visual Camera 0 Visual Camera 1.5 THERMAL THERMAL 0 5 CONTROL CONTROL Total 25 Total 17 Margin -5 % Margin +29 %

 Positive margin for long science phase. Slight negative margin during insertion orbit. Thermal Environment

Q Absorbed Heat Load(W/m2) Q incident (W/m2) Emittance Absorptivity (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

Q Absorbed Heat Load(W/m2) Q incident (W/m2) Emittance Absorptivity (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

   Qabsorbed=αε() QQQ solar ++ Mars Phobos + Q MarsIR 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 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 Generous Support ////