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.

LOGIC Mission Concept

Trajectory 7 Month Mars Transfer Orbit

Highly Elliptical Capture Orbit

→

Trajectory Phobos Intercept

→

6-8 Months Aerobraking

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.

Spacecraft

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

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, 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

Generous Support

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