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Microsatellite Design and Integration for INSPIRE: International Satellite Project in Research and Education
Loren Chang, Upper Air Dynamics Lab
Institute of Space Science, National Central University, Taiwan
2016-10-25 APSCO & ISSI-BJ Space Science School
Microsatellite Design and Integration for INSPIRE: International Satellite Project in Research and Education
Loren Chang, Upper Air Dynamics Lab
Institute of Space Science, National Central University, Taiwan
2016-10-25 APSCO & ISSI-BJ Space Science School
Nano
C.K. Chao, Jude Salinas, Ya-Chih Mao, Jack Chieh Wang, Jia-Yu Su, Duann Yi, Pei-Yun Chiu, Joe Hong, Yi-Chung Chiu Institute of Space Science, National Central University, Taiwan Steven C.R. Chen National Space Organization, Taiwan Amal Chandran, Michael McGrath Laboratory for Atmospheric and Space Physics, University of Colorado, USA Dave Fritts, Larry Gordley GATS, USA Priyadarshan, Kastubh Anand Indian Institute of Space Science and Technology, India Tarun Pant Space Physics Laboratory, ISRO, India
INSPIRESat team
Overview
• The INSPIRE Project
• Designing a Mission Concept: • Scientific Needs: The Middle and Upper Atmosphere
• Science Traceability: Doppler Wind and Temperature Sounder (DWTS)
• Can a satellite do this? Feasibility, System Definition, and Lessons Learned
Space Science and Capacity Building at NCU
INternational Satellite Program In Research and Education
• CubeSat constellation for Earth and Space Science Research (2 missions currently
planned)
• Global ground station network
• Academic program in mission & spacecraft design, operations, and data analysis.
University of
Colorado LASP
Indian Institute of Space
Science and Technology /
ISRO
Taiwan National Central
University / NSPO
INSPIRESat-1 Feasibility Study Subsystem NCU CU IIST
Payload
Science
STR
TCS
CDH
COM
EPS
ADCS
Launch
http://www.theonion.com/article/nasa-announces-plan-to-launch-700-million-into-spa-1950
Why is this mission needed?
Mission Concept
Review (MCR)
To evaluate the feasibility of
the proposed mission
concept(s) and its fulfillment
of the program's needs and
objectives.
To determine whether the
maturity of the concept and
associated planning are
sufficient to begin Phase A.
http://nodis3.gsfc.nasa.gov/displayDir.cfm?Internal_ID=N_PR_7120_005E_&page_name=Chapter2
TROPOSPHERE
STRATOSPHERE
MESOSPHERE
THERMOSPHERE
You are here
Earth’s Atmosphere
~ 20 km
~ 60 km
~ 85 km
Drag is the largest orbital perturbation in LEO!
Orbital debris distribution
https://orbitaldebris.jsc.nasa.gov
GEO
MEO
LEO
TROPOSPHERE
STRATOSPHERE
MESOSPHERE
THERMOSPHERE
You are here
Satellite Drag
Reentry Dynamics
Ionospheric Variability
Communications
Stratospheric Polar
Vortex
Weather Prediction
Middle & Upper Atmosphere (>20 km)
~ 20 km
~ 60 km
~ 85 km
TROPOSPHERE
STRATOSPHERE
MESOSPHERE
THERMOSPHERE
You are here
~ 20 km
~ 60 km
~ 85 km
Limited Observational Ability
Gordley et al., SPIE 2013
Doppler Wind and Temperature Sounder (DWTS)
• New gas filter correlation radiometer measurements of wind and temperature (~25 – 250 km)
• Cross-track limb pointing IR camera with cryocooler.
• Designed and patented by GATS. Fabrication by University of Colorado LASP.
• Still requires operational heritage in space environment (TRL 4).
Limb scan of NO emissions
Doppler shift and width of
emission lines
Atmospheric winds and
temperature from 30 to 200
km
DWTS Design & Interface
15
draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum
electrical power draw of the cooler of 20W, thus we have significant margins to drive the cooler
harder. Thermal control is maintained through a combination of active and passive methods.
The optics (including the gas cell) are maintained to < 240K. For the flight instrument, the heat
pulled by the cooler is transferred through heat pipes using non-toxic fluids to radiators on side
panels. Remaining thermal control is through heaters and passive conduction/radiation. The
envisioned constellation of “free-flyers” will allow much cooler optics and gas cells, which
reduces noise and narrows the cell line-widths improving the DWTS DIP signature.
Figure 13. Dark current density versus
temperature for InAsSb photodiode and nBn
detector, and HgCdTe photodiodes with a cut-off
wavelength of 5 μm. From: Performance
comparison of barrier detectors and HgCdTe
photodiodes (Opt. Eng. 2014;53(10):106105.
doi:10.1117/1.OE.53.10.106105)
OPTICS: The draft DWTS optical design is shown in Figure 14. This design supports the
SB204 detector form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal
length, 30 mm aperture, resulting in a fast F/1.4 system. The lens includes mixture of
Germanium and ZnSe lenses as required for stray light control and image quality. Both
spherical and aspherical elements are used, as the optics shall be diamond turned to the required
surface figure and finish.
Figure 14. DWTS Optical Design
ELECTRONICS: The camera electronics shall be 4-5 Watts from previously built boards.
The cryocooler controller electrical budget is included in the 10W given. Thus the power for the
instrument needed shall be 14-15 Watts (without reserves) or 23 Watts with 50% reserves. The
mass is estimated to be less than 4 Kg (1 kg for the Integrated Detector Cooler Assembly, .75
Kg for the mounted optics, 0.75 Kg for the electronics and cable hardnesses, and 1.5 Kg for the
structure and mounts. A radiation tolerant by software design approach shall be taken using
Figure 15. Single Channel DWTS Sensor Head
14
and IDL save sets. Useful mission documents, data user guide, published papers, and software
tools (usually in IDL) for reading and plotting the DWTS data will be provided on a DWTS
project web site on CU/LASP web server. The LISIRD databases are designed for large (TB)
data sets, so all of the DWTS data will be kept on-line, and the LISIRD backup system. We
intend to preserve public access of the data for the life of the LISIRD system. However, LISIRD
is not configured, nor funded, to be a long-duration archive center, so DWTS archive will be at
a NASA archive site.
2.2. DWTS Instrument Design
The DWTS instrument design leverages existing
technologies and commercial off-the-shelf (COTS) hardware.
The heart of the instrument is the Lockheed-Martin Santa
Barbara Focalplane SBF204 nBn Focal Plane Array with a
Ricor K527 cryocooler, integrated into a Brandywine micro-
cryocooled camera. Figure 12. Ricor K527 Split
Linear Cryocooler
DETECTOR: The detector format is 1296 x 1040 (8 pixel borders) with a 12-micron pitch
as shown in Table 2. This FPA is produced in large volumes (>50/month) for tactical programs,
with similar SBF nBn FPAs having passed radiation testing at the Air Force Research
Laboratory. nBn was chose over HgCdTe due to having COTS FPAs in stock, and having better
epitaxial growth properties (better pixel gain uniformity and higher operability.)
Parameter SBF204
Format: 1296 x 1040 (8 pixel borders)
Windowing: Windowing: 608 x 8 up to 1280 x 1024 in 1 x 4 pixel increments.
Pixel Size: 12 µm x 12 µm
Detectors: nBn with QE = 40% @ 5.25 um
Outputs: Selectable 1, 2, or 4 LVDS Data pairs, plus 1 LVDS clock pair
Integration Modes: 1: Snapshot, Integrate Then Read
2: Snapshot, Integrate While Read
Line Rate: < 125 kHz in four port mode
Frame Rate: < 99 Hz for 1280 x 1024 in four port & 14 bit mode
A/D Resolution Selectable, 14 or 13 bits
Signal Capacity: 50 Ke-
Total ROIC Noise: < 30 e-
Responsivity: 3 e-/LSB
Dark Current <10 nA/cm2 @110K (from graph for 5um cutoff)
Power: < 250 mW
Table 2. Lockheed Martin SBF204 Specifications
THERMAL CONTROL SYSTEM: Our operating requirements are a heat lift capacity of
400 mW, cold tip temperature of <90K while at a rejection temperature of -20K. The
manufacturer has confirmed that we can easily meet these requirements with an electrical power
Subsystem Notes
Camera Teledyne GEOSNAP FPA (in fabrication)
2000 x 2000 pixels
Ricor K527 cryocooler
Optics Includes NO gas cell. NO 5.3 𝜇m (MidIR)
absorption.
43.3 mm effective focal length
30 mm aperture.
7° x 20° FOV
Cooler & Heater Gas cell temperature control.
Processor Xiophos Inc. Q6 FPGA
Data aggregation.
0.5 Hz sample return to bus: FPA data,
cooler and heater settings.
Data storage and power provided by bus.
Test Objectives
• Mission shall demonstrate ability of DWTS to
measure winds and temperatures from 30 –
200 km.
• DWTS shall be demonstrated in a space
environment.
• Demonstrate that stray light effects can be
eliminated using optical or digital means,
including lunar scanning to demonstrate stray
light calibration.
• The satellite shall resolve planetary-scale
waves and tides with zonal wave numbers 1
to 4 from the lower to the upper atmosphere
from 60S to 60N.
Mission Objectives (Initial)
• The mission shall survive the space environment.
• The mission life-time shall be at least 3 months.
• The data rate shall be 225 megabits/day.
• The communication system for the mission shall be VHF and UHF radio frequency bands.
Customer Top Level Requirements
• The mission shall have four ground-stations- Colorado, Taiwan, Singapore and India.
• The mission shall be launched by ISRO as a secondary payload.
• The spacecraft shall be disposed of at the end of life.
• The satellite bus shall be a CubeSat with 6U dimensions.
Will a 6U CubeSat be able to support the DWTS
payload and mission?
Determine payload requirements.
Set subsystem requirements that
satisfy payload requirements.
Identify available COTS options that
satisfy payload requirements.
*Feasibility Study funded by National Space
Organization of Taiwan for SY 2015-2016
System
Requirements
Review (SRR)
To evaluate whether the
functional and performance
requirements defined for the
system are responsive to
the program's requirements
on the project and represent
achievable capabilities.
http://nodis3.gsfc.nasa.gov/displayDir.cfm?Internal_ID=N_PR_7120_005E_&page_name=Chapter2
DWTS Requirements & Specifications
Mass 4 kg
Volume 29 x 10 x 9 cm (3U)
Power Standby and Operational: 7 W
Safehold: 1 W
Off: 0 W
Pointing
Knowledge
±0.5 arcmin, all axes
Pointing Stability < 6 arcsec/sec
Attitude Control 1° all axes
Field of View Sun most not appear in field of
view.
Spacecraft Velocity
Knowledge
±1 m/s
Data Rate Downlink 200 Mbits/day
Thermal Stability FPA: ±0.1 K/minute
Gas Cell: ±1 K/minute
Thermal
Requirement
Anti-Sunward Side: -10 – 0°C
Observed Emission NO 5.3 𝜇m (MidIR)
15
draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum
electrical power draw of the cooler of 20W, thus we have significant margins to drive the cooler
harder. Thermal control is maintained through a combination of active and passive methods.
The optics (including the gas cell) are maintained to < 240K. For the flight instrument, the heat
pulled by the cooler is transferred through heat pipes using non-toxic fluids to radiators on side
panels. Remaining thermal control is through heaters and passive conduction/radiation. The
envisioned constellation of “free-flyers” will allow much cooler optics and gas cells, which
reduces noise and narrows the cell line-widths improving the DWTS DIP signature.
Figure 13. Dark current density versus
temperature for InAsSb photodiode and nBn
detector, and HgCdTe photodiodes with a cut-off
wavelength of 5 μm. From: Performance
comparison of barrier detectors and HgCdTe
photodiodes (Opt. Eng. 2014;53(10):106105.
doi:10.1117/1.OE.53.10.106105)
OPTICS: The draft DWTS optical design is shown in Figure 14. This design supports the
SB204 detector form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal
length, 30 mm aperture, resulting in a fast F/1.4 system. The lens includes mixture of
Germanium and ZnSe lenses as required for stray light control and image quality. Both
spherical and aspherical elements are used, as the optics shall be diamond turned to the required
surface figure and finish.
Figure 14. DWTS Optical Design
ELECTRONICS: The camera electronics shall be 4-5 Watts from previously built boards.
The cryocooler controller electrical budget is included in the 10W given. Thus the power for the
instrument needed shall be 14-15 Watts (without reserves) or 23 Watts with 50% reserves. The
mass is estimated to be less than 4 Kg (1 kg for the Integrated Detector Cooler Assembly, .75
Kg for the mounted optics, 0.75 Kg for the electronics and cable hardnesses, and 1.5 Kg for the
structure and mounts. A radiation tolerant by software design approach shall be taken using
Figure 15. Single Channel DWTS Sensor Head
15
draw of 10 Watts steady state, and 15 Watts peak for 3-minutes. This is half the maximum
electrical power draw of the cooler of 20W, thus we have significant margins to drive the cooler
harder. Thermal control is maintained through a combination of active and passive methods.
The optics (including the gas cell) are maintained to < 240K. For the flight instrument, the heat
pulled by the cooler is transferred through heat pipes using non-toxic fluids to radiators on side
panels. Remaining thermal control is through heaters and passive conduction/radiation. The
envisioned constellation of “free-flyers” will allow much cooler optics and gas cells, which
reduces noise and narrows the cell line-widths improving the DWTS DIP signature.
Figure 13. Dark current density versus
temperature for InAsSb photodiode and nBn
detector, and HgCdTe photodiodes with a cut-off
wavelength of 5 μm. From: Performance
comparison of barrier detectors and HgCdTe
photodiodes (Opt. Eng. 2014;53(10):106105.
doi:10.1117/1.OE.53.10.106105)
OPTICS: The draft DWTS optical design is shown in Figure 14. This design supports the
SB204 detector form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal
length, 30 mm aperture, resulting in a fast F/1.4 system. The lens includes mixture of
Germanium and ZnSe lenses as required for stray light control and image quality. Both
spherical and aspherical elements are used, as the optics shall be diamond turned to the required
surface figure and finish.
Figure 14. DWTS Optical Design
ELECTRONICS: The camera electronics shall be 4-5 Watts from previously built boards.
The cryocooler controller electrical budget is included in the 10W given. Thus the power for the
instrument needed shall be 14-15 Watts (without reserves) or 23 Watts with 50% reserves. The
mass is estimated to be less than 4 Kg (1 kg for the Integrated Detector Cooler Assembly, .75
Kg for the mounted optics, 0.75 Kg for the electronics and cable hardnesses, and 1.5 Kg for the
structure and mounts. A radiation tolerant by software design approach shall be taken using
Figure 15. Single Channel DWTS Sensor Head
SUB-
SYSTEM
MASS
PAYLOAD 4.000 kg
ORB/NAV 0.012 kg (pqNAV)
EPS 0.580 kg (GOMSpace
20 SA panels)
CDH 0.088 kg (Pumpkin)
COMM 0.078 kg (Astrodev)
0.100 kg (Antenna)
ADCS 0.850 kg (XACT)
STR 1.640 kg (6U Chassis)
TOTAL 7.348 kg
Margin 0.652 kg
Mass Budget
ORBIT, GUIDANCE & NAVIGATION (ORB/NAV) ATTITUDE DETERMINATION AND CONTROL SUBSYSTEM (ADCS) ORB/NAV Manager: Jack Chieh Wang
ADCS Manager: Cornelius Csar Jude H. Salinas
ORB/NAV Requirement: The mission orbit shall be at least
350 km in altitude and be at near-circular orbit.
NASA Debris
Assessment
Software
650 km: 37.7 years
800 km: >100 years
Debris Mitigation
Deorbit Guidelines:
25 years
ORB/NAV Requirement: The mission orbit shall be at least
60N in inclination.
At 60 N, local time precession
rate is:
39.7097 days at 650 km altitude
42.0544 days at 800 km altitude
ORB/NAV Requirement: The mission shall monitor the
orbit via ground-tracking and GPS.
COTS Option for Navigation
Best Option
pqNAV-L1/FM
(SkyFox LabsTM )
SGR-05U
(Surrey)
GPS Receiver
(SSBV)
GPSRM 1 GPS Receiver Module
(CubeSat Kit)
Feature
GPS L1 C/A
signal, 15
channels
GPS L1 C/A
signal,
12 channels
GPS L1 C/A
signal,
12 channels
GPS L1/L2/L2C signal,
GLONASS L1/L2 signal
Price 4900 EUR =~5360
USD ? ? ?
Mass 12 gram 40 gram < 30 gram ?
Dimensions 42*42*15 mm 70x45x10mm 50x20x5mm ?
Power 120 mW
(3.3V@25°C) 800 mW(5V) < 1W (5V) ?
Temporal
resolution 1 Hz 2x106 Hz > 1 Hz ?
Operation
Temperatur
e
-40~+85°C -20~+50°C -10~+50°C ?
Altitude Up to 3600 km ? (Already flown) ? (Already flown) ?
Velocity Up to 9 km/s ? (Already flown) ? (Already flown) ?
ADCS
REQUIREMENTS
ADCS CAPABILITY
BC XACT
• Payload shall always point
cross-track and away from
the sun.
• Attitude knowledge: +/- 0.5
arcminutes
• Attitude control: +/- 1
degree for all direction
angles.
• Pointing stability: <6.0
arcsecs/sec.
• Attitude knowledge: +/-
0.06 arcminutes for all
direction angles
• Attitude control:
+/- 0.003 degree (1-sigma)
for 2 axes
+/- 0.007 degree (1-
sigma) for 3rd axis
• Pointing stability:
1 arcsec/sec for all axes
XACT: On orbit tests on MinXSS
http://digitalcommons.usu.edu/smallsat/2016/S5GuidCont/2/
THERMAL CONTROL SUBSYSTEM (TCS) TCS Manager: Sunny Duann-Yi
Components Operational Survival
EPS -40 ~ 85 ℃ -50 ~ 90 ℃
Batteries 5 ~ 20 ℃ 0~ 25 ℃
Solar Panels -150 ~ 110 ℃ -200 ~ 130 ℃
DWTS Payload -30 ~ 20 ℃ -35 ~ 25 ℃
GPS -40 ~ 85 ℃ -55 ~ 95 ℃
Reaction Wheels (ADCS) -10 ~ 40 ℃ -20 ~ 50 ℃
Motherboard -40 ~ 85 ℃ -45 ~ 90 ℃
Antennas -100 ~ 100 ℃ -120 ~ 120 ℃
Worst Case Estimates: Radiative Equilibrium
-69.4 – 50.1 C at 600 km
-72.2 – 48.8 C at 800 km
Lesson Learned: TCS
Payload requirements should be frozen before a
preliminary design for the spacecraft.
14
The DWTS optical design is shown in Figure 13. This design supports the GEOSNAP detector
form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal length, 30 mm
aperture, resulting in a fast F/1.4 system. The lens includes mixture of Germanium and ZnSe
lenses as required for stray light control and image quality. Both spherical and aspherical
elements are used, as the optics shall be diamond turned to the required surface figure and
finish. The optics will have low-emissivity anti-reflection coatings to keep internal emissions
low. The inside surface of the lens barrel and the sensor housing will be coated black to reduce
reflected stray light.
Thermal Design: The thermal design has two actively-controlled thermal zones, the detector,
which must be kept at 95 K (110K is acceptable) to reduce dark current to our goal of
approximately 100 e-/pixel/s, and the cryocooler radiator which rejects heat at 23 C. In
addition, the cryocooler will use passive cooling to keep the optics 210 K to reduce thermal
emission. A Ricor K527 Stirling split-linear cryocooler will be used to cool the detector. The
K527 has a cooling capacity of 0.7W for an active
load at 95K and requires 20W input power. A
flexible thermal strap will be used to connect the cryo
cold head to the detector, providing vibration
isolation from the cryocooler pistons. A large
deployable radiator will be used to reject the heat
generated by the cryocooler. Precise control of the
temperature in both zones will be accomplished via
heaters placed on the detector housing and on the
radiator. The detector will be mounted to the lens
assembly using thermally-isolating mounts, and then
both are enclosed inside a double walled thermal
enclosure. The lens assembly will be mounted to the
enclosure also using thermal isolating mounts. To maintain radiative isolation, the outside
surface of the lens barrel and the inside surface of the enclosure walls will be gold coated.
Figure 12 shows the instrument concept thermal model. End of Life (EOL) material properties
were used for hot case analysis and Beginning of Life (BOL) properties were used for cold case
analysis. In the worst case hot orbit scenario, the cryocooler active load is 0.54 W, which has
30% margin to the K527 capacity of 0.7W. Additional margin could potentially be gained
through detailed design of the isolating mounts and increased detector flex cable length. The
maximum predicted heater power required to maintain the detector and radiator at the design
temperatures are 0.25W and 16W respectively. There are also maneuvers that can be
performed, such as 180 degree role about the line of sight (effectively flipping the FOV, which
does not affect the measurements), that will dramatically improve radiator and solar array
efficiency. This strategy was not fully evaluated before the proposal, but will conservatively
add 30% to our cooling margins.
Figure 12. Thermal Model of Sensor and Platform.
• Later payload component level thermal analysis by
LASP engineers found that a radiator is needed for
sufficient heat dissipation from DWTS cryocooler.
• Cryocooler introduces 60 Hz disturbance torque that
must be characterized by duty cycling cryocooler
during payload operation.
• TCS and ADCS requirements require revision to
accommodate.
COMMUNICATIONS (COM)
COM Managers: Ken Lai, Pei-Yun Chiu, Kitty Jia-Yu Su
HOUSEKEEPING DATA: 1. Orbital position and
velocity (ORB) 2. Battery Condition (EPS) 3. Power Consumption
(EPS) 4. Sub-system and
payload current temperature (TCS)
5. Attitude (ADCS) 6. Angular Velocity (ADCS) 7. Universal time 8. Error Feedback
RAW MISSION DATA: 2D Radiance in Pixel Values
FINAL MISSION DATA: Temperature and
Horizontal Wind Vertical Profiles
GROUND RECEIVING STATION 1. Mission Control
2. Ground-based processing
USERS: Researchers, companies and mission center
new commands
COMMUNCATIONS ARCHITECTURE:
The space-craft will utilize a centralized communication system. All data-processing will be done on the ground.
STK Analysis at altitude 800 km and inclination 65N
Total access time/day in
worst case
Colorado, U.S 6,250 Seconds
SARC,
Singapore 3,286 Seconds
ISRO, India 3,444 Seconds
NCU, Taiwan 2,400 Seconds
Total 15,380 Seconds
With the 225 megabits per day requirement, it is required that the
communication module onboard the spacecraft must achieve a minimum
data rate of 14.629 kbps.
COM
REQUIREMENTS
COM DESIGN
• COMM data rate: 225
megabits/day.
• Transmit to four ground
stations: Colorado, Taiwan,
Singapore and India.
• Data rate: 590 megabits per day
• Ground stations: Colorado,
Taiwan, Singapore and India
• COMM Module data rate:
> 38.4 kbps
• VHF and UHF radio frequency
bands COMM system.
• AX.25 packet protocol
• GMSK Modulation downlink.
• COMM Module Frequency
bands: VHF and UHF
• Protocol: AX.25 Packet Protocol
• Modulation: GMSK Modulation
AstroDev Helium 100 Radio
38.4 kbs
(120 – 150, 400 – 450 MHz)
ISIS Deployable Antenna
NCU Ground Station • VHF/UHF Amateur
radio bands.
• HyGain 218SAT antennas with dual bandpass filters.
• ICOM 9100 transceiver.
• Kamtronics KAM-XL terminal node connector.
Lesson Learned: COM
Remember to account for reduced tracking speeds during high
elevation angle passes, and losses at low elevation angles.
Typhoon Magi
NCU Weather Station
Vmax = 54 km h-1
Lesson Learned: COM
Exercise caution in defining ground operations SOP and
be strict on implementation!
ELECTRICAL POWER SUBSYSTEM (EPS) EPS Manager: Cornelius Csar Jude H. Salinas
TLR/PAYLOAD
REQUIREMENTS
EPS REQUIREMENTS EPS DESIGN
1. Science goal TLR
2. Engineering goal TLR
3. 3-month Mission Lifetime
4. DWTS power requirement is 7W.
5. DWTS has four modes: Off (0W),
Safehold (1W), Standby (7W) and
Operational (7W)
• Adequate power for at least 3
months.
• Total power: 17.13 W (0.8565 W
margin)
• Solar panel power : 44 W
• Battery capacity: 20 Wh
• Protection against: Bus Voltage,
Bus Faults, Under Voltage, Over
Voltage, Over Current and Single
Event Effects.
• Switching mechanism such that
the switching is on the positive
side.
• Protection Mechanisms: Under
voltage, over voltage (3.3 V, 5 V
and 12 V) and over current
protection. Temperature sensor.
EPS micro-controller unit.
SUB-
SYSTEM
COTS OPTION WITH HIGHEST POWER
REQUIREMENT
PAYLOAD 7 W (DWTS)
ORB/NAV 1.3 W (GPSRM)
EPS -
CDH 0.003 W
COMM 6 W (Astrodev Helium Radio)
ADCS 2.83 W (XACT)
TOTAL 17.133 W(maximum)
EPS Requirement: The EPS shall provide for a required
power consumption of 17.133 W (0.8565 W margin).
InSPIRESat-1
MISSION ORBITAL PERIOD 100 minutes
ECLIPSE PERIOD
(SPAT Analysis)
1.9758 – 35.0728 minutes
TOTAL POWER
CONSUMPTION
17.13 W
REQUIRED SOLAR
PANEL POWER
(assuming 30%
efficiency power loss for
all transfers)
25.1709 W to 43.1479 W
EPS Requirement: The solar panel shall provide 44 W of
power for the spacecraft.
InSPIRESat-1
MISSION ORBITAL PERIOD 100 minutes
ECLIPSE PERIOD
(SPAT Analysis)
1.9758 – 35.0728 minutes
TOTAL POWER
CONSUMPTION
17.13 W
REQUIRED BATTERY
CAPACITY (assuming
30% efficiency and 80%
DOD)
17.8809 Wh
EPS Requirement: The EPS battery shall have a required
capacity of at least 20 Wh.
Are there COTS solar panels that can provide 44 W?
GOMSpace NanoPower P110 Series solar panels (http://gomspace.com/?p=products-p110).
Clyde Space 3U Solar Panel (https://www.clyde.space/products/25-3u-solar-panel).
Pumpkin Space Systems 3U Solar Panel (http://www.cubesatkit.com/docs/press/P
umpkin_CSDWLU_2010-2.pdf).
OPTION 2:
3U Size at
7S1P
Configuration
7W at BOL
OPTION 3:
3U Size at
8S1P
Configuration
8W at BOL
OPTION 1:
60.36 cm2 Size
1 Solar Panel
2.3 W at EOL
ESTIMATED NO. OF
PANELS AND AREA:
20 panels. 1207.2 cm2.
Are there COTS batteries that can provide 20 Wh?
Clyde Space 20 Whr stand-alone battery
(https://www.clyde.space/products/22-20whr-
cubesat-battery).
Clyde Space 30 Whr stand-alone battery
(https://www.clyde.space/products/9-30whr-
cubesat-battery).
Clyde Space 40 Whr stand-alone battery
(https://www.clyde.space/products/49-40whr-
cubesat-battery).
GOMSpace battery pack 38.9 Wh (http://gomspace.com/document
s/ds/gs-ds-batteries.pdf).
OPTION 1:
20 Wh capacity
OPTION 2:
30 Wh capacity
OPTION 3:
40 Wh capacity
OPTION 4:
40 Wh capacity
Are there COTS power-distribution modules with
built-in voltage and current protector?
GOMSpace P31us power distribution module (http://gomspace.com/index.php?p=products-p31us).
Clyde Space FlexU power distribution module (https://www.clyde.space/products/19-3rd-generation-flexu-eps).
Blue Canyon Tech XEPS electrical power sub-system module (http://bluecanyontech.com/wp-content/uploads/2015/05/EPS-Data-Sheet_1.0.pdf).
OPTION 1:
3 Solar Panel Inputs for
maximum of 30 W.
Voltage: 3.3 V and 5 V.
OPTION 2:
3U to 12U deployable solar
panels
Voltage: 3.3 V, 5 V and 12 V
OPTION 3:
6 Solar Panel Inputs for maximum of 60 W.
Voltage: 3.3 V, 5 V and 12 V. 20Wh/40Wh battery pack.
COTS Options for Solar Panel Configuration
ESTIMATED NO. OF
PANELS AND AREA
FOR 44 W:
20 panels. 1207.2 cm2.
Solar Panel Configuration Constraints:
1. DWTS shall point cross-
track.
2. DWTS shall avoid pointing
directly at the sun.
PROBLEM: Zero flux at 0 degree beta-angle.
Solar Panel Configuration Constraints:
1. DWTS shall point cross-
track.
2. DWTS shall avoid pointing
directly at the sun.
SOLUTION 1: Solar panel drives.
Solar Panel Configuration Constraints:
1. DWTS shall point cross-
track.
2. DWTS shall avoid pointing
directly at the sun.
SOLUTION 2: Additional orthogonal solar panels.
Solar Panel Configuration Constraints:
1. DWTS shall point cross-
track.
2. DWTS shall avoid pointing
directly at the sun.
SOLUTION 3: Periodic charging mode. Reduced payload duty cycle
(off during eclipse).
SOLUTION 1 SOLUTION 2 SOLUTION 3 1. Added mass.
2. Additional costs.
3. Additional attitude
control.
4. Additional electronics.
5. Driver malfunctioning.
1. Added mass.
2. Additional costs.
3. Additional attitude
control.
4. Deployment risk.
1. Frequent slewing.
2. Reduced data
coverage.
Test Objectives
• Mission shall demonstrate ability of DWTS to
measure winds and temperatures from 30 –
200 km.
• DWTS shall be demonstrated in a space
environment.
• Demonstrate that stray light effects can be
eliminated using optical or digital means,
including lunar scanning to demonstrate stray
light calibration.
• The satellite shall resolve planetary-scale
waves and tides with zonal wave numbers 1
to 4 from the lower to the upper atmosphere
from 60S to 60N.
Mission Objectives (Revised)
Will encounter 0 solar flux every 90° of orbital
precession.
Total power budget, solar panel requirement and battery capacity
requirements are estimated. COTS options for space-grade and
prototype solar panel, battery and power distribution modules are
identified.
Options on solar panel configurations that can support the
operational test of a single-channel DWTS are identified.
Mission objectives must be revised such that they can be
achieved at the expense of reduced data coverage.
Will a 6U CubeSat be able to support the DWTS
payload and mission?
Risks Identified
Subsystem EPS ADCS Thermal
Risk High power requirements
No Sun at certain beta angles
Attitude knowledge and control
requirements
Expected temperature range exceeds
component survival range.
Mitigation
Options
• Charging mode with reduced
payload duty cycle
• Steerable solar panels
• Perpendicular fixed arrays
• Blue Canyon XACT ADCS
module
• DWTS cryocooler duty cycling
to characterize disturbance
torque smearing.
• DWTS radiator pointed at deep space.
• Component level thermal modelling.
• Passive thermal control components
for non-payload subsystems.
Steady State Temperature Range (600
km):
-69.4C ~ 50.1C
Components Operational
EPS -40 ~ 85C
Batteries 5 ~ 20C
Solar Panels -150 ~ 110C
DWTS -30 ~ 20C
GPS -40 ~ 85C
ADCS -10 ~ 40C
Motherboard -40 ~ 85C
Antennas -100 ~ 100C
DWTS Radiator -20 ~ 0C
�
� = � �°
� = �°
�
Pointing
Knowledge ±0.5 arcmin,
all axes
Pointing
Stability < 6
arcsec/sec
Attitude
Control 1° all axes
Feasibility & Iteration Revised mission objectives can be
achieved with a 6U CubeSat at the
expense of reduced data coverage.
Payload design must be revised, with
laboratory testing of in-orbit calibration
methods to reduce risk.
Additional details:
Chang, L.C. et al. (2016), A Preliminary Design for the
INSPIRESat-1 Mission and Satellite Bus: Exploring the
Middle and Upper Atmosphere with CubeSats,
Proceedings of the AIAA/USU Conference on Small
Satellites, LEO Missions, SSC16-WK-02.
http://digitalcommons.usu.edu/smallsat/2016/S4LEOMis
/5/.
14
The DWTS optical design is shown in Figure 13. This design supports the GEOSNAP detector
form factor and pixel pitch, using a 7.0°x 20.0° FOV, 43.3 mm effective focal length, 30 mm
aperture, resulting in a fast F/1.4 system. The lens includes mixture of Germanium and ZnSe
lenses as required for stray light control and image quality. Both spherical and aspherical
elements are used, as the optics shall be diamond turned to the required surface figure and
finish. The optics will have low-emissivity anti-reflection coatings to keep internal emissions
low. The inside surface of the lens barrel and the sensor housing will be coated black to reduce
reflected stray light.
Thermal Design: The thermal design has two actively-controlled thermal zones, the detector,
which must be kept at 95 K (110K is acceptable) to reduce dark current to our goal of
approximately 100 e-/pixel/s, and the cryocooler radiator which rejects heat at 23 C. In
addition, the cryocooler will use passive cooling to keep the optics 210 K to reduce thermal
emission. A Ricor K527 Stirling split-linear cryocooler will be used to cool the detector. The
K527 has a cooling capacity of 0.7W for an active
load at 95K and requires 20W input power. A
flexible thermal strap will be used to connect the cryo
cold head to the detector, providing vibration
isolation from the cryocooler pistons. A large
deployable radiator will be used to reject the heat
generated by the cryocooler. Precise control of the
temperature in both zones will be accomplished via
heaters placed on the detector housing and on the
radiator. The detector will be mounted to the lens
assembly using thermally-isolating mounts, and then
both are enclosed inside a double walled thermal
enclosure. The lens assembly will be mounted to the
enclosure also using thermal isolating mounts. To maintain radiative isolation, the outside
surface of the lens barrel and the inside surface of the enclosure walls will be gold coated.
Figure 12 shows the instrument concept thermal model. End of Life (EOL) material properties
were used for hot case analysis and Beginning of Life (BOL) properties were used for cold case
analysis. In the worst case hot orbit scenario, the cryocooler active load is 0.54 W, which has
30% margin to the K527 capacity of 0.7W. Additional margin could potentially be gained
through detailed design of the isolating mounts and increased detector flex cable length. The
maximum predicted heater power required to maintain the detector and radiator at the design
temperatures are 0.25W and 16W respectively. There are also maneuvers that can be
performed, such as 180 degree role about the line of sight (effectively flipping the FOV, which
does not affect the measurements), that will dramatically improve radiator and solar array
efficiency. This strategy was not fully evaluated before the proposal, but will conservatively
add 30% to our cooling margins.
Figure 12. Thermal Model of Sensor and Platform.
Revised Mission Timeline
19
3.2. Project Schedule
The DWTS top-level schedule is shown in Figure 16. The project will start on January 21,
2017 on receipt of funds. The project will be part of the graduate course in Spacecraft Design in
Spring 2017. The schedule balances design and review efforts with the need to order critical,
long lead, items early enough to preserve a robust system I&T effort. The Preliminary Design
Review (PDR) will be completed, seven months after program start on August 20, 2017 and
The Critical Design Review (CDR) is scheduled five months after PDR. Since many subsystems
are commercial off the shelf (COTS) parts and the XB1 spacecraft bus comes in a defined
package, the procurement of long lead parts (such as solar arrays, XB1, detector from Teledyne)
begins after PDR.
Figure 16. DWTS Top-Level Schedule
Three months are allocated to mechanically integrate system, make any necessary
mechanical adjustments, and perform initial telemetry
checks and verify subsystem handshaking. Following
system integration, two months are allocated to vet
operational test scripts, perform a CPT, and perform an
end-to-end test from the MOC to the spacecraft through the
ground test equipment. Environment tests proceed after
baseline compatibility and performance testing is complete
on the integrated flight hardware with two
3.3. Risk Reduction Plan
Because of their limited on-board resources, modest
development budgets, and commitment to involving
relatively inexperienced student teams in their design and
implementation, CubeSat projects present a higher level of
risk than would be acceptable for most space flight
missions. However, the overall approach to continuous risk
management used by larger programs applies equally to
small ones: identification of risks, categorization by
likelihood and consequences, planning of mitigation strategies, and tracking of progress
Color Code. Green=acceptable w/o
further mitigation; Yello=
mitigation necessary, acceptable at
launch; Red=risk drivers, aggressive
mitigation need, mitigate before launch.
Additional Lessons Learned (So far)
• For a payload defined mission, freezing payload requirements is essential before the spacecraft bus can proceed to PDR levels.
• Project management and standard operating procedures are essential and should be strictly defined, especially for international collaborations.
• Good communication and design iteration is essential between the science and engineering teams.
• Never underestimate the value of good PR for securing support and team building!
INSPIRESat-2 / Compact Ionosphere Probe
• Schedule: • Payload and bus PDR: 12/2016
• Payload and bus CDR: 8/2017
• Funded by Taiwan MOST, University of Colorado.
STR
Requirements
● All the structures shell not cause Interference.
● Mass budget shell be less than 4kg, and mass margin is +2kg.
Assembly and Interior layout
3U CubeSat, size is 474 x 1113 x 337mm with deployed antennas. Total mass is
2.598kg without harnesses, screws and GPS antenna. It is not possible to reduce
the size from 3U to 2U because the mass will overload for 2U structure. CubesatAIP
has total 2855 parts.
CDH
CDH Requirements
18
INSPIRESat-2: Subsystems Breakdown Subsystem NCU CU IIST
Payload
STR
TCS
CDH
COM
EPS
ADCS
Software
Launch
Thank you. Questions?
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