Devils Logic PDR presentation

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INTERPLANETARY CUBESAT DESIGN
IRA FULTON SCHOOL OF ENGINEERING
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Science
Question
Science
Objective
Physical
Parameters
Observables
Surface
composition of
Phobos
Determine
Mineral /
Chemical
composition
Spectral
Reflectance
in IR range
Spectral Reflecatnce sampling
at 10 nm bandwidths from
1000nm to 2400nm
Historical
nature of
Phobos' surface
structure and
morphology
Imaging of
striation / crater
intersection
points
Structure
and
morphology
High resolution images of
striation intersection points
Location of L1
Stability
Observe the
stability of the L1
Lagrange Point.
Position &
Velocity of
CubeSat
 Spacecraft position
 Doppler velocity
 Spacecraft Attitude
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Science Objectives
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Science Mission Overview
• Achieve orbit around Mars beyond Phobos
• Perform multiple fly-bys of Phobos
• Begin use of primary instruments
o Infrared Spectrometer
o High resolution visible light camera
o Collect Star Tracker and IMU data
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Improvements Upon Existing
Science Data
• Spectroscopy
• Surface reflectance spectra will improve data previously
collected by
o Mars Pathfinder
o Mariner 9
o Viking Lander
• This information will be used to
o More accurately classify the moons’ composition
o Improve understanding of the origin of Phobos
• High-Resolution Visible-light Imaging
• Visible light surface images collected by LOGIC will improve
upon data collected by
o Mars Express HRSC Imager
 50 kilometer altitude
 5 meters/pixel resolution
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Instrument Challenges
Spectrometer
• Decreasing movement between spacecraft and camera
during image capture to avoid motion blur
• Filtering out data which exceeds transmission data
budget
Visible-light Camera
• Ensuring that the surface is within the field of vision
for a variety of orbital altitudes
• Lens must remain undamaged and free of debris
ADCS
• Must reject erroneous data
• High-accuracy clock for precise signal-delay
measurements
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Principle Mission Challenges
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Science Traceability Matrix
Science
Question
Science
Objective
Physical
Parameters
Observables Instruments
Required
Instrument
Performance
Projected
Instrument
Performance
Mission
Requirements
Surface
composition
of Phobos/
Stickney
crater
Determine
Mineral /
Chemical
composition
Surface
reflectivity
in IR range
Spectral
reflectance
sampling at
6 nm
bandwidths
from
1000nm to
2400nm
Spectrometer
·<300 m/pixel
·Numerous
narrow spectral
bands
·Low integration
time
·<100 m/pixel
·>100 spectral
bands (<7nm)
·<4.096 s
integration time
·<38 km Altitude
above Phobos
·Attain orbit
around Mars
above Phobos
Historical
nature of
Phobos'
surface
structure
and
morphology
Imaging of
striation /
crater
intersection
points
Structure
and
morphology
High
resolution
images of
striation
intersection
points
High-
Resolution
Visible Light
Camera
·<5 m/pixel res.
·Large FOV
·65-500
mm/pixel res.
·14.25° FOV
·<14 km Altitude
above Phobos
·Attain orbit
around Mars
above Phobos
Gravity field
near Phobos
Measure
gravitational
field
strength
Mass
distribution
/volume of
Phobos
·Spacecraft
position
·Doppler
velocity
·Spacecraft
Attitude
·Radio
·Deep Space
Network
·ADCS
·<1m/pixel
(0.01°) accuracy
·>1°/s
·ADCS
Accuracy>1E4
·<1m/pixel
(0.01°) accuracy
·>3.5°/s
·ADCS
Accuracy>1E8
Frequent comm-
checks between
LOGIC and
Earth
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Science Objective Instrument Minimum requirement
Determine Mineral /
Chemical
composition
Spectrometer
 Spectral range 1000 nm to 2400nm
 10 nm Spectral band
 100 channels
 100m x 100m spatial resolution
Imaging of striation /
crater intersection
points
Visual
Spectrum
Camera
 Minimum Resolution of 1.8 m / pixel
 1/3rd imaging of surface
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Requirements & Challenges
Challenges:
• Volume Constraint: under 10%
• Low Albedo: 7.1% (2.31 Lux)
• Low data relay rates: 7.8 kbps
• Spectral range selection :(VIS-NIR-SWIR)
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Spectrometer
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Spectrometer
Parameter
Edmond Optics
1000 2000nm
InGaAs NIR
NIR Quest 256-
2.5
Argus 1000
Mass [g] 650 1180 230
Volume [cc] 1020.6 940.7 180
Power [W] 12 15 6.2
Range [nm] 1000 - 2000 900 - 2500 1000 - 2400
Spectral Resolution
[nm]
8 9.5 12
No of bands 128 128 100
No of pixels 256 Pixel Array 256 Pixel Array 256 Pixel Array
Integration Time 20 µs to 10 s 1-400 ms 0.5µs to 4.1s
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0
0.5
1
1.5
2
2.5
3
3.5
4
4.5
EDMOND OPTICS NIR QUEST ARGUS 1000
Edmond Optics NIR Quest Argus 1000
Power consumption[Whr] 3 3.75 1.55
Data per exposure[kb] 4.096 4.096 3.328
Spectrometer Power Consumption & Data Generation
Power consumption[Whr] Data per exposure[kb]
Spectrometer
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Visual Spectrum Camera
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Cameras
Parameters
CIRES - E2V Malin ECAM-C50
Teledyne Dalsa
Genie
Mass [g]
Sensor system 70 256 196
Lens system 240 100 460
Volume [cc]
Sensor system 50.92 199.1 129.7
Lens system 103.5 269.7 367.2
Power [W] 1.5 2.5 4.5
Pixel Density [MP] 1.3 5 12
Fly-by
SR [m/pixel] 1.8 1.8 1.8
Max WD [m] 6912.00 14310.00 22118.40
Nom Case
SR [m/pixel] 2083.33 1006.29 651.04
WD [m] 8000.00 8000.00 8000.00
Best Case
SR [m/pixel] 1302.08 628.93 406.90
WD [m] 5000.00 5000.00 5000.00
WD: Working distance SR: Spatial resolution
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0
5
10
15
20
25
30
35
40
CIRES - E2V MALIN ECAM-C50 TELEDYNE DALSA GENIE
CIRES - E2V Malin ECAM-C50 Teledyne Dalsa Genie
Power Consumption[Whr] 0.375 0.625 2.375
size of single image[MB] 3.75 14.74 36
Camera Power Generation and Data Generation
Power Consumption[Whr] size of single image[MB]
Visual Spectrum Camera
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Recommendations
Camera
description
Weight
age
Spectrometer Visual Spectrum Camera
Edmond
Optics
NIR Quest
Argus
1000
CIRES -
E2V
Malin
ECAM-C50
Teledyne
Dalsa
Genie
Mass 15% 3 1 5 5 4 3
Volume 15% 1 2 5 5 3 4
Power &
Operating
temperature
10% 2 1 3 5 4 3
Spectral
resolution
15% 4 3 1 2 4 5
Performance 45% 4 4.56 4.33 4.22 3.67 1.78
Space heritage Aerospace Aerospace Yes Leo Yes No
Total 100% 64% 61% 78% 84% 74% 58%
The performance is a function of spectral range , integration time , SNR & QE and image size for spectral and lens
mass & vol , shutter speed , SNR& QE and image size for the camera
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Architecture
Spectrometer
Visual Spectrum camera
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Modifications
1. Improvement of spectral resolution: Larger pixel array
2. Improvement of spectral range : different diffraction grating
3. Reduction of the volume : use commercially available lens
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Single CubeSat Limitations
• Volume constraint : Volume availability for science payload
is limited (less than 1U)
• Power constraint : Unable to operate science payloads and
communication system simultaneously
• Low data transmission
o Small antenna and low power
o Nominal transmission rate (7.8kbps)
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Benefits of CubeSat Network
• Volume
o Better quality science instrument
o Multiple Spectrometers to cover the required spectrum
• Power
o Reduction in power requirement
o Simultaneous subsystem operation capability
• Data transmission rate
o Extra power for transmitter
o Larger antennas
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www.nasa.gov
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Maximization of Science Data
• Spectrometer
o Cover complete spectrum: UV,VIS, IR & Xray
o Investigation for more mineral compositions
o Improved mineral classification capability
• Visual camera
o More information and improved interpretation capability
o Stereo vision-lead to 3D map
o Enhancement of features
• Stability Data
o Inter-CubeSat communication, S band ranging
o Improved accuracy
o Analogues to GRAIL and GRACE
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CubeSat Network Vs Hayabusa I
Instrument Hayabusa I CubeSat Network
Spectrometer
NIR: range (700-2100 nm)
XRF:0.7 - IO KeV
Spectrometer network
can obtain more spectral
data (362-3920nm)
Visual Spectral
Camera
Multiband Came
Resolution -70cm at 7km
Resolution -62cm at 5km
LIDAR
Range: 50m-50km
1-m resolution
-
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Cubesat Network Architecture
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Cubesat Modifications
Cubesat I & Cubesat II
• Visual spectrum camera on both CubeSats
• UV-VIS spectrometer on CubeSat I
• IR spectrometer on CubeSat II
• S-band antenna and transponder for inter-CubeSat
communication and data transmission to relay CubeSats
Cubesat III & communication relay sat
• X-ray spectrometer on CubeSat III
• No science payload on communication relay sat
• S-band antenna and transponder for inter CubeSat
communication
• X band antenna and transponder for communication with Earth
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1. Collect a representative
sample of spectral data
from Phobos’ with 100
bands and 300m
resolution
2. Image 1/3rd of Phobos’
surface at a minimum
resolution of 1.8 m/ pixel
3. Observe the stability of
the L1 Lagrange Point
Mission Objectives
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• 6U CubeSat
• Spectral and Visible Sensor Array
• Deployable Solar Panels
oHawk MMG Gimballed Deployable
• Deployable X-Band Antenna
oISARA JPL
• Dual Propulsion Systems
oAerojet Green Monopropellant
oBusek Electrospray
System Architecture
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www.planetarysystemscorp.com/
http://www.mmadesignllc.com/
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System Design
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• Deploy from Launch Vehicle
o Systems check, telemetry check and initial burn
• Cruise Phase (208 Days)
o Idle payload and propulsion with limited communication
• Mars Capture (270 days)
o 22 min impulsive burn over 90 min to reduce 870 m/s
o Aero-braking for 135 days to steadily circularize orbit
o EP thrust for 135 days to reach Home orbit ~200km from
Phobos
Concept of Operations
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Mission Operations (547 days)
◦ Weekly schedule allotting 4 hour window of DSN
communication per week
◦ Depart Home orbit and approach Phobos to get data
◦ Collect 5 visible and 5 spectral images within 15 km
◦ Return to Home orbit and transmit data
◦ Transmit 5 spectral and 5 thumbnails of visible
images
◦ Select best 2 thumbnails and transmit cropped
lossless visible image data
Concept of Operations
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Budgets / Feasibility
Mass, Volume and Power Budget
Subsystem
Mass [kg]
(Max 14)
Volume [cc]
(Max 7000)
Power [W]
(Max Capture 44)
Chassis 1.000 7000
Power 2.110 700 0.7
Communication 2.440 508 12.9
ADCS 0.850 500 3.0
Propulsion 3.897 3080 15.0
Payload 0.586 591 2.65
Thermal 0.061 60 0.5
Total / Margin 10.944 / 21.8% 5439 / 32.0% 34.75 / 22.8%
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Programmatic Risks
Key Challenges
• Test failures
• Quality rejections
• On time delivery
• Cost variations
• Supplier availability
• Mission obsolescence
Risks and Challenges
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• Unable to capture into Mars orbit
oDe-scope tertiary objective and accomplish both primary
and secondary objectives with a fly-by
• Unable to achieve mass/volume budget
o De-scope secondary objective and accomplish both primary
and tertiary
o Remove Malin eCam-C50 from Payload
o Reduction in mass of 0.356 kg – Improves Margin by 4%
o Reduction in volume of 411 cc – Improves Margin by 5%
De-scope Considerations
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Top Level Requirements
• Compliance with NASA 6U CubeSat standards
• Spectral data in range of 1000 nm to 2400 nm at 100m x 100 m
spatial resolution
• Capture 33.3% of Phobos surface at 1.8 m pixel resolution
• Comply with the NASA General Environmental Verification
Standard
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Instrument Down
Selection Approach
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• Science Data Collection
• Data transfer rate
• DSN availability
• Unexpected Communication black outs
• Transit Time (7 months ) and time to achieve phobos orbit
(9 months)
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Parameters Affecting
Mission Duration
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Instrument Down Selection
Approach
Communications
Science
Time Best Case Nominal Case
Best
(Co-orbit @ 5 km)
Data transfer 12 months 32 months
Mission time 28 months 48 months
Nominal
(Co-orbit @ 8 km)
Data transfer 5 months 12 months
Mission time 21 months 28 months
Worst
(Fly By @ 14 km)
Data transfer 1.5 months 3.6 months
Mission time 17.5 months 19.6 months
Mission time varies between 17.5 to 28 months based on above conditions
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Risk Assessment
 Vibrations
 EMI interferences
 Radiation
 Temperature Variations
 Space Debris Impact
 Hardware failures such
as bit error ,chip error
 Software Malfunctions
 Outgassing of material
 Degradation of Material
Strength
 Effect on the cube
sat reliability
 Inaccurate of science
data
 Reduction in Mission
Life
 Complete Mission
Failure
 Use of off the shelf
components which
have good space
heritage
 Redundant
subsystems
 Allocating task to
alternative
subsystem
 Ground testing of
software and
hardware
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Risk Assessment - Design
Complete Mission
Failure
Crtical Reduction of
Misision Life leading
to reduction in the
performance of the
components
Reduction in
Accuracy of
science data
Effects the
performance of the
other subsystems
leading to mission
failure in the long
run
Frequent
(Highest Probability of
occurrence )
Radiation effecting
onboard Comps
Unexpected short
duration
communication losses
Moderate
ADCS pointing
inaccuracy due to
External EMI
Solar Panel Failure due
to external impact or
bending loads
Faulty orientation
of antenna
Antenna
Deployment/Solar
panel deployment
Failure due to gimbal
failure
Occasional Propulsion System
Ignition Failure
1.External EMI
interfernce with EPS
and Controllers
2. Active Thermal
system Malfunction
Components
Outgassing
leading camera
lens fogging
Reduction in bolted
joint pretension due
to Creep
Remote
Structural failures due
to Vibration 2. Impact
by Space Debris
Camera Startup
failure
Uneven thermal
Expansion of
structure
Severity of
Risk
Risk
Occurrence
Probability
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System Configuration
Comparison
Subsystem Configuration-1 Configuration-2 Configuration-3
Battery
Capacity
>120 W h 120 W h 120 W h
Antenna X– Band X-Band X-Band & UHF
Payload
Camera & Point
Spectrometer
Only Spectrometer
Point Spectrometer &
Camera
Pros and Cons
• Longer
Communications
• Lots of Science
Data
• Basic Model
• No High
Resolution
Cameras
• Better
Communications
during orbit
insertion
• Lots of science
data
• More mass and
less data
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System Requirements
Computer
● Plug and play architecture
● Space heritage
● Low active mode power consumption
● USART, SPI, I2C interfacing support
● Health check and autonomous fault
repair
● Signature check algorithm
● Use of CCDS standards
Telecom
● X-band communication
● DSN compatibility (<-190 dB)
● EIRP > 22 dB
● Turnaround ratio (880/749)
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Components
Iris X-band Transponder
http://www.clyde-
space.com/cubesat_shop/obdh/
pumpkin_cubesat_obc/pumpkin_motherboard
MSP 430
Pumpkin Motherboard
https://store.ti.com/cc3100boost-
cc31xxemuboost-exp430f5529lp.aspx
https://store.ti.com/cc3100boost-
cc31xxemuboost-exp430f5529lp.aspx
High Gain Reflectarray
http://mstl.atl.calpoly.edu/~bklofas/Pres
entations/DevelopersWorkshop2015/Kle
sh_MarCO.pdf
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Computer System
Architecture
MSP430
SD Card
4 GB
Voltage
Regulator
RAM
512 B
ADC
10-bit
Clock
16 MHz
Flash
16 KB
Watchdog
15-bit
5 V
I2C
To EPS
USCI
(SPI and
USART)
To
Transponder/
ADCS/
Propulsion
To other
subsystems
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Functional Flow and
Pseudo Code
1. Establish
Communication Link
3. Control trajectory
2. Get T&T and
science data
1. Main system
check
3. Power system
check
2. Housekeeping
system check
4. Communication
link check
6. Science data
capture
5. Trajectory
check
Pseudo CodeFFBD
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Antenna Analysis
Efficiency Gain (dB)
Number of
deployable
Gain (dB)
55% 30.35 1 31.69
75% 31.69 2 34.7
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Link and Data Budget
Table 1 : Storage requirements (in MB)
Table 2 : Achievable data rates (in bps)
Figure 1: Data accumulation plot for best-best case
Comms
Science
Best Worst
Best 766.48 3802.9
Nominal 309.54 1503.5
Worst 103.18 471.68
Spectrometer 0.332
Range
Gain
Closest Nominal Farthest
Best 19679.43 9782.92 3885.43
Nominal 12302.78 7815.12 2700.46
Worst 6860.08 4942.85 901.76
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Results
● Hardware and software
requirements have been fulfilled
● Communication link can be
established
● 200 MB of data margin
Results and
Recommendations
Recommendations
● Use of Ka band
● Use DSN for longer time
● Get better EIRP
● Use of LDPC code
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Propulsion & Mission Trajectory
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Planetary Capture
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∆𝑉 = 𝑉∞
2 + 2
𝜇
𝑅 𝑚 + 𝑎𝑙𝑡
− 2
𝜇
𝑅 𝑚 + 𝑎𝑙𝑡
−
𝜇
𝑎
𝑎 =
𝑅 𝑚 + 𝑎𝑙𝑡 + 𝑅 𝑆𝑂𝐼
2
Lowest Energy Capture
Periapsis : 120 km altitude
Apoapsis : 576,000 km
Eccentricity ~ 1
Capture ∆V
Time of flight : 207 days
V∞ = 2/438 km/s
∆Vcap = 569.2 m/s
Aerobraking ∆V = 4.3 m/s
© Devils Cube

Capture Propulsion System Trade Study
Planetary Capture
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Propulsion System Specific
Impulse
Propellant
Mass
Fraction
Propellant
Mass
Propellant
Volume
HYDROS
Bipropellant
300 s 0.176 1.794 kg 1794 cm3
Busek Green
Monopropellant
220 s 0.232 2.365 kg 1577 cm3
Aerojet Green
Monopropellant
250s 0.212 2.112 kg 1405 cm3
© Devils Cube

Capture Propulsion System Trade Study
Planetary Capture
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Propulsion System Propellant
Mass
Propellant
Volume
Thrust System
Mass
System
Volume
HYDROS
Bipropellant
1.794 kg 1794
cm3
0.8 N 1.2 kg 1000 cm3
Busek Green
Monopropellant
2.365 kg 1577
cm3
0.5 N 1.5 kg 500 cm3
Aerojet Green
Monopropellant
2.112 kg 1405
cm3
4 N 1.3 kg 1100 cm3
© Devils Cube

Planetary Capture
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Thrusting Schedule
◦ ∆Vcap = 569.2 𝑚
𝑠
◦ 𝑚 = 0.0016 𝑘𝑔
𝑠
◦ Thrust = 4 N
◦ Optimal Maneuver at 120km
Required Thrusting Time
◦ 22 Minutes*
◦ Finite Burns Increase ∆V
◦ 1.5 Hours < 9000 km
◦ ∆Vmax = 870 𝑚
𝑠 +40%
◦ 4 Hours < 23000 km
◦ ∆Vmax = 1190 𝑚
𝑠 +80%
◦ Thermal Dependence
© Devils Cube

Initial Conditions
• Periapsis : 120km Altitude
• Apoapsis : 576,000 km
Procedure
◦ Raise Periapsis to ≈ 150km
◦ LOGIC passes through mars atmosphere
If Periapsis < 120 km
Raise Periapsis to ≈ 150km
Elseif Apoapsis < Phobos
Raise Periapsis to ≈ 300km
End Aerobraking
Aerobraking
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Aerobraking
∆V: 350 m/s
Approx Time : 4.5 Months
© Devils Cube

Initial Conditions:
• Periapsis : 300 km Altitude
• Apoapsis : 9300 km
Final Conditions:
• Periapsis : 9234 km
• Apoapsis : 9517 km
Phobos Co-Orbit
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Impulsive ∆V: 285 𝑚
𝑠
Low Thrust ≈ 385 𝑚
𝑠
Low Thrust Transfer Time:
≈ 4.5 months
Propulsion
System
Specific
Impulse
Propellant
Mass
Fraction
Propellant
Mass
Propellant
Volume
System
Volume
Busek
Electrospray
2300 s 0.0177 0.175 kg 175 cm3 300 cm3
Busek Pulsed
Plasma
536 s 0.071 0.703kg 231 cm3 330 cm3
Aerojet Green
Monopropellant
250 s 0.109 1.08 kg 720 cm3 N/A
© Devils Cube

Mission Trajectory
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• Orbit around Mars while staying
close to Phobos for long time
• Requirements
o Not crash into Phobos
o Stay in close range
• Optimal Condition
o Stay as close as possible for
long time to make maneuvers to
perform science mission
Home Orbit
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Orbital Parameter
Semi-major axis 9378 km
Eccentricity 0.0073
Inclination 1.22 degree
Min. Distance 35.1 km
Max. Distance 233.1 km
Results from Analysis with STK
• Inclination is relative to Mars’s Equatorial Plane
• Inclination of Phobos is 1.075 deg
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Distance to Phobos surface in a day
Orbit path in Phobos’ fixed frame
In a day
Orbit path in phobos’ fixed frame
In a month without maneuvers
© Devils Cube

ADCS Trade Study
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Product Pointing
Accuracy
Slew
Rate*
Volume Mass Power
XACT ± 0.007 deg 4.3 deg/s 10 x 10 x 5 cm 0.85 kg 2W (steady)
3W (max)
MAI-400 ± 0.05 deg 2.7 deg/s 10 x 10 x 5.59
cm
0.635 kg 4W (steady)
8.5W (max)
MAI-200 < ± 0.05 deg 2.7 deg/s 10 x 10 x 7.87
cm
0.907 kg 5.5W(steady)
13.7 W (max)
XACT has the best performance in
almost all categories.
*Moment of inertia was estimated as 0.2 kg-m2 for 6U Cube Sat to calculate the slew rate
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On the Transfer Orbit to Mars
o No Camera Pointing is required
o Antenna Pointing to Earth
(4 hours per week)
o Solar Panel Orientation to Sun
(as much as possible)
o Thrust Vector Control
(as reaching to Mars)
Pointing Schedule
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Pointing Schedule on the
Transfer Orbit to Mars
© Devils Cube

Co-Orbit with Phobos
• Communication
4 hours every week
• Power Charging
Whenever possible, the maximum
surface of solar panel should be
facing to the Sun.
Pointing Schedule
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Angle profiles to the important
directions from the direction vector
of CubeSat on the orbit in a day
© Devils Cube

Considerations
63
Considerations for Orbit
• Maneuver Schedule
oTo cover wide range of the surface
• Delta V consumption
o In order to perform maneuvers to collect
enough science data
Considerations for Pointing
• External torque
o Solar pressure
o Gravitational torque
o Torque from rotating solar panels
• Momentum damping requirement
o Use of thruster for damping
© Devils Cube

Keep Alive Power Requirements
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Power Budget for Keep Alive Configuration
System Subsystem Power Survival (W) Power Active (W)
Comms Antenna (X-Band) 0 0
Transponder 1 1
CPU/OBC 0.01 0.01
Subtotal 1.01 1.01
Power Photovoltaics 0 0
Battery 0.1 0.2
EPS 0.1 0.1
Subtotal 0.2 0.3
Thermal MLI 0 0
Heaters 5 1
Subtotal 5 1
ADCS Sealed Unit 0.85 1.5
Subtotal 0.85 1.5
Propulsion Electrospray 0.5 5
Thruster 0.5 0.5
Subtotal 1 5.5
Payload Camera + Lens 1.75 2.5
Spectrometer 0.9 6.3
Subtotal 2.65 8.8
Total 10.71 18.11
© Devils Cube

Power Schedule
66
• Communications are ON for 4 hours, once a week (Peak Power)
• Thermal, constant at maximum
• Payload 3x a day, for 30 minutes duration each time
• Propulsion 1x per day, for one hour
• ADCS working at constant operational level
© Devils Cube

Power Architecture
67
• Peak Power Tracking:
o Longer mission durations
o Solar array can be decoupled
(simper array designs)
• Centralized Architecture:
o Distributes all voltages rails
from one location
o Fewer regulators required
© Devils Cube

Solar Insolation Model
68
Date May 7 2021 Jun 2 2021 Feb 7 2022
Eclipse
time
(average)
38 min 25 min 53 min
Duration of eclipses for orbit next to Phobos
© Devils Cube

Power System State
69
• Li-ion batteries (120 Whr) – DoD 31.9% :
o 1% discharge per day, 90% EPS efficiency
o Communications are ON for 4 hours, once a week
o Thermal, constant at maximum (5 W)
o Payload 3x a day, for 30 minutes duration each time (2.65 W idle/8.8 W peak)
o Propulsion 1x per day, for one hour (1 W idle/5 W active)
o ADCS working at constant operational level (1.5 W)
© Devils Cube

Energy Capability of the Batteries
Component Model Voltage (V) Capacity (Ah) Energy (Wh)
Battery
GS NanoPower
BPX
Up to 29.6 V 2.6 154
Power Configurations
71
Energy Available to Charge the Batteries
Component Model Efficiency Energy (W)
Deployable Solar
Panels
Hawk Solar Arrays -
MMA
30% 22
Electrical Power System
Component Part Number Efficiency
Power Consumption
(W)
EPS Blue Canyon Tech EPS 85% <0.1
o Li-ion batteries
o Autonomous heater system
o Can be configured for nominal voltages ranging up to 29.6 V
o Sun tracked continuous high power
o 140 W/Kg
o Modular and scalable to 100 W peak power and 50 W
OAP
o Charge and distribution fault protection
o Space heritage
© Devils Cube

• Ensure System is warm and temperature doesn’t drop below
253K
• Ensure subsystems temperature is within the operating limits
• Transient Thermal Environment
• Radiation is the main mode of heat transfer
Objectives & Challenges
73© Devils Cube INTERPLANETARY CUBESAT DESIGN
© Devils Cube

Transient Thermal
Environment
• Variations in Solar flux, Mars IR, over a number of days considered
• Exact values of AU found using Wolframalpha
• Eclipse period during one orbit found using STK
Spacecraft Thermal Control Handbook Volume 1: Fundamental Technologies David G. Gilmore
74
Perihelion Aphelion Mean
Direct Solar (W/m2) 717 493 589
Mars Albedo 0.29 0.29 0.29
Maximum IR (W/m2) 470 315 390
Minimum IR (W/m2) 30 30 30
AU 1.381 1.666 1.5235
© Devils Cube

Heat
Load(W/m2)
Q incident
(W/m2) Emittance Absorptivity
Q Absorbed
(W/m2)
Solar Flux 638.9407 0.55 0.35 223.629245
Mars Albedo 191.6822 0.55 0.35 67.08877
Phobos Albedo 44.7259 0.55 0.35 15.654065
Mars IR 425.9474 0.55 0.35 234.27107
Total Maximum 540.64315
75
Heat
Load(W/m2)
Q incident
(W/m2) Emittance Absorptivity
Q Absorbed
(W/m2)
Solar Flux 492.6368 0.55 0.35 172.42288
Mars Albedo 147.4846 0.55 0.35 51.61961
Phobos Albedo 34.4846 0.55 0.35 12.06961
Mars IR 315.5439 0.55 0.35 173.549145
Total Minimum 409.661245
Transient Thermal
Environment
© Devils Cube

• Thermal modelling done using ANSYS steady-state
thermal and transient thermal
• Assumptions:
o Approximate geometric shapes
o Ignored effects of mountings
o Ignored effect of solar panels and antennas
o Geometry scaled to 1/4th actual size
o Actual battery dimensions
Thermal Modelling
77
© Devils Cube

From Earth to Mars
• 207 days required to travel
from LEO to Mars orbit
• Steady state analysis
• Patch heaters (0.5 W) & ADCS
switched on
79
© Devils Cube

Critical Subsystem
Battery
• Operating temperature in
the range of -15o C to 75o C
• Prolonged exposure to
freezing temperatures affect
charge transport and cause
electrode damage
• Steady state simulations
performed
80
© Devils Cube

Thermal Sub-System
Requirements
81
Characteristic Description Characteristic Description
External MLI Single Aluminized
Kapton (eight
layers of 1 mill. )
Heaters Kapton
Resistance Patch
Heaters
Emissivity
Absorptivity
0.55
0.35-0.51
Quantity Four
Max. Weight 11.4 g Power 0.5 W each
Internal MLI Single Aluminized
Kapton (One
Layer of 0.5 mill.)
Controlling
Mechanism
Tayco Solid State
Controller
Emissivity
Absorptivity
0.03
0.14
Density 0.0019 g/cm2
Component to be
covered
Propulsion
system(Except
Nozzle)
© Devils Cube

Mitigation Strategies
• MLI has very good space heritage
• Heaters at risk of failure
• Turn on batteries to maintain temperature in case of
heaters failure
82
© Devils Cube

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