dino pdr 7 october 2015 deployment and intelligent nanosat operations university of colorado at...
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DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
University of Colorado at Boulder
University Nanosat III
August 14th-15th, 2003Logan, Utah
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Industry Support
We would like to thank the following companies who have donated hardware and mentorship to Colorado Space Grant Students working on the DINO program at the University of Colorado at Boulder. Without their support, DINO would never leave the ground.
PLANETARYSYSTEMSCORPORATION
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DINO
The purpose of the student-led Deployment and Intelligent Nanosatellite
Operations (DINO) mission is to determine cloud heights from space,
evaluate the performance of intelligent operations, and assess deployment
technologies for nanosatellites including a tether, memory composite hinges, and
thin-film solar arrays.
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Objectives
• Science and Education– Stereoscopic imaging to determine cloud heights– Student leadership and training– Active involvement of K-12 Students
• Deployment demonstrations– Thin-film solar array deployment– Gravity-gradient tether deployment– A simple deployment using memory composite hinges
• Intelligence – Onboard evaluation of science and engineering data
– Autonomous response and rescheduling to optimize ops
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Overall Mission TimelineLaunch
ICU DeploymentDINO Deployment
Mission Activation
Tether Deployment
FITS Deployment
Aerofin Deployment
Normal Science Operations
Detumble
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Spacecraft Level RequirementsRequirement Method Status
The spacecraft must not exceed a mass of 30 kg. Design, Test
The spacecraft must operate on 30W or less. Design, Test
The spacecraft’s center of gravity (CG) shall be within 0.25” of the geometric central axis of the ICU.
Design, Analysis
The allowable static envelope of the spacecraft is a cylindrical right prism with a diameter of 18.7” (47.5 cm) and a height of 18.7” (47.5 cm).
Design
The spacecraft’s CG shall not lie more than 12” above the satellite interface plane (SIP) .
Design, Analysis
The spacecraft shall have a fundamental frequency above 100 Hz given a fixed-base condition at the SIP.
Design, Analysis
The spacecraft must be capable of meeting all mission objectives.
Design, Test
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Experiments and Mission Highlights
• Deployables– Gravity Gradient Tether (GGT)
• Tip Mass Communications
– Foldable Integrated Thin-film Solar Arrays (FITS)– Elastic Memory Composite (EMC) Hinges
• Stereoscopic Imaging• Intelligent Operations• Thin-film Solar Arrays and Lithium Polymer
Batteries
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1. Gravity Gradient Tether and Tip MassRequirement Method Status
Tip Mass must have a mass of 5 kg. Design, Test
Tether must be greater than 10 m. Design
Tip mass must stay within 60º from zenith to retain gravity gradient stability during deployment and mission lifetime.
Design, Analysis
The Tip Mass tip-off rate should be less than 3 º/s. Analysis
Complete deployment must occur in sunlight to avoid thermal snaps. Analysis
The Tip Mass must have its own power supply and meet all NASA safety requirements for batteries.
Design
The Tip Mass must be able to capture a picture after each deployment and send them to the main spacecraft.
Design, Test
The Tip Mass deployment systems must have 10W @ 12V for 1 minute. Design, Test
The Tip Mass must not recoil at the end of the deployment. Design, Analysis
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1. Deployment Method
• Open-Loop Deployment– Lightband will provide kickoff velocity of 2 ft/s
• Deployment will take approximately 40 sec
– Tether will be “left-behind” by tip mass– Braking system will slow tip-mass near end of travel– Simple compared to a complex motor system
Main Satellite
Braking System
Tether Z-fold
Tip Mass
Lightband
Tether Guides
Velocity
Tether
Wheel (turning)
Brake shoe (fixed)
Brake
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1. Donated Material
• Tether– Low-density polyethylene tape
• Along its length are three strands of 0.003 inch thick Spectra® 1000
– 0.005 inch by 1 inch cross-section– Heritage
• Flown on Advanced Tether Experiment (ATEx) mission
• Lightband– 15 inch motorized separation system– Delta V = 2 ft/s– m ≈ 6.5 lbm– Tip-off rate < 1º/s– Flight proven
Tether Material
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1. Tip Mass• Main Mission
– Provide gravity-gradient stabilization for DINO
• Experiment Mission– Take and send pictures of the
following deployments:• Tether• FITS• Aero-fins
• Success Criteria– A picture is taken after each
deployable.– These pictures are
successfully sent to the main satellite and then to Mission Operations.
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1. Tip Mass Features
• Structure• Imaging
– Camera trade study ongoing
• Communications– Uses wireless LAN card on
tip mass and main module
• Power– One Lithium-polymer
(same as main module)– 5V DCDC converter
CameraEPS
Tether BoxComm
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1. Tip Mass Communication
• Standard Wi-Fi LAN card • Wiser 2400 interfaces wireless 802.11 link with RS232
serial link from the camera
Processor with a PCMCI slot for Wi-Fi card
Digital camera
802.11 Wi-Fi card
802.11 air interface
Wiser 2400 unit (OTC wireless)
RS-232 serial interface port of WISER 2400
Serial RS232 port of camera
802.11 interface port of WISER 2400
Main sat module Tip Mass
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2. FITS and Release Requirements
Requirement Method Status
The FITS system must be preloaded to 100 lbs. Analysis
The release mechanism for the FITS system must have a holding force of 200 lbs or greater.
Design, Test
EPS must provide a 28V line for each mechanism. Design
A complete side panel must be available for each FITS system and its release mechanism.
Design
EPS must provide 25W for 30 sec for each release mechanism.
Analysis, Test
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2. FITS System - Stowed
Deployment Hinge
Restraint Panel
Separation Device
• Provided by Microsat
• Released with Frangibolts
• Preloaded to 100 lb
• Upon release deployment is almost instantaneous
• Stowed Envelope
• 12.5 x 7.25 x 1.3 in
• Volume = 0.067 ft3/Wing
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2. Frangibolt
Release Execution• Memory Composite • 500 lbs holding force• 25W @ 28V • 21 seconds• Increased temperature
from power activates release.
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2. FITS Deployment
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3. EMC Hinges and Release Mechanism
Requirement Method Status
4 separation switches must be monitored by C&DH to ensure full deployment.
Design, Test
The composite panel aerofins cannot bear any loads. Analysis, Test
EPS must provide a 28V bus for the EMC hinges and the release mechanism.
Design
EPS must provide 18W for 2 minutes for the release mechanism. Analysis, Test
EPS must provide 10W for 1 minute for each EMC hinge. Analysis, Test
Each pair of hinges must controlled thermally to maintain a temperature between 88°-92°C.
Design, Test
A side panel must be kept open for each aerofin and its mechanism.
Design, Test
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3. Aerofins
Elastic Memory Composite Hinges
Provided by CTD
10W @ 28V for 1 min
Provided by Starsys
High Output Paraffin (HOP) Actuator
18W @ 28V for 2 min
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3. Aerofins Release Mechanism
StowedReleased
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4. Science Requirements
Requirement Method StatusThe science system must not exceed a mass of 0.56 kg. Design,
Test
The system must operate on less than 11W while taking a picture. Design, Test
The camera used must have a field of view between 40-55 degrees. Test, Analysis
The camera must have sufficient resolution to determine cloud heights within 500m.
Test,
Analysis
The shutter speed of the camera must be 1/60th of a second or faster. Design
All components shall comply with NASA’s outgassing specifications Analysis
Any glass components shall comply with NASA’s regulations Test, Analysis
All components shall meet NASA’s low-released mass part Analysis
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4. Imaging Summary• Purpose
– The primary mission of DINO is stereoscopic imaging of cloud formations.
– Overlapping images of the same object from different angles provides a 3-dimensional photograph.
• Design– Two cameras aimed at +/- 30°
along track• AIPTEK Pen Cam 1.3 Mega• 41.5° FOV; 1248 x 960 max res
– Software converts into topographic maps
• Success Criteria– A topographic map is sent to the
ground and has been verified as correct.
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4. Science Mission
• Analysis of images will be performed onboard the spacecraft.
• Images processed and ranked for coloring and edge distinction.
• Image opportunities assessed from analysis of first image.
• Stereo images combined to form a topographic map of cloud heights for downlink.
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5. Intelligent Operations RequirementsRequirement Method Status
The on-board data system should support the migration of autonomy through tailoring of flight applications software and updates to procedures, rules, constraints, and others.
Design, Test
Operational constraints, rules, target synchronization, timeline, and algorithms shall be updated by users and autonomously as the situation allows.
Design, Test
Schedules should be automatically generated from commands users select on a website.
Design, Test
The selection of a ground station to communicate with the satellite should be chosen autonomously.
Analysis, Test
Demonstrate of rescheduling a task with improved performance. Design, Analysis, Test
Determine if an image is a good opportunity for a stereoscopic image and then change the schedule to take another picture.
Design, Analysis
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5. Intelligent Operations
• Virtual Mission Operations Control Center (VMOCC) will provide automated web-based spacecraft control.
• Spacecraft Command Language (SCL) will provide onboard schedule execution and fault detection and reaction.
• Real-time spacecraft data will be available via the internet for educational outreach.
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5. Intelligent Operations
• VMOCC automated capabilities– Create a schedule for the satellite from users worldwide– Input orbit event data from STK into schedule– Automatically initialize a radio (anywhere in the world) and send
the command– Accept telemetry from any radio worldwide and integrate into
VMOCC model
• SCL capabilities– Assess image opportunities– Execute scheduled commands in real-time– Perform onboard fault detection and reaction– Maintain database for sensors, spacecraft states, and hardware
performance.
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6. Electrical Power SystemRequirement Method Status
EPS must not exceed a mass of 2.25 kg Design, Test
The system must be able to accept commands and execute those commands from C&DH through an RS-232 port.
Design, Test
Battery cells will last the lifetime of the mission or longer which is approximately 6000 orbits.
Design, Analysis
EPS will provide a 5V, +/- 12V and 28V buses for all the subsystems to use. Analysis
The operational temperature for the batteries must be kept between 0o and 40oC. Analysis, Test
Satellite will be un-powered while on the Shuttle. Design, Test
The lithium battery system will be two fault tolerant. Design
Each cell will have individual temperature monitoring for the satellite and GSE. Design
All inhibits must be able to have status checked without the satellite powered from the GSE.
Design, Test
3 independent inhibits must be used for the Lithium Polymer batteries. Design
The FITS system must be able to provide an average of 30W of power over an entire orbit for the lifetime of the mission.
Design, Analysis
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6. Electrical Power System
EPS Board
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6. Thin-Film Solar Array Summary
Beginning of Life End of Life
25° Celsius78° Celsius – Radiation
Degradation
Area 1.10 m2 1.10 m2
Voltage – String 17.4 Volts 13.0 Volts
Current – String 1.4 Amps 1.27 Amps
Power – Total Watts 85 Watts 64 Watts
Array Efficiency 5.70% 4.60%
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6. Lithium Polymer Battery Design• Characteristics:
– 3.7 V potential, 4 A-hr capacity, 120 g each
– Non-flammable– Prismatic internal structure
• Safety– 2-fault tolerant – Cell vents will not be oriented downward
at any time during launch on the shuttle– Cells will have thermistors for
temperature monitoring – Charging/discharging
• The cells will be maintained above 3.2V
• The stack will be nominally charged to 95% of capacity to prevent overcharge of individual cells
– Battery assembly will be maintained between 0° and 40°C
Credit: Valence Electronicswww.valence.com
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System Functional Diagram
Imaging
Cam A Cam B
Legend
USB
ADCSTR1
TR2
TR3
MAG
Sun Sensors
x6
Rate Gyro x3
Data Bus
Power Lines
Thermal
+5V Line
+/-12V Line
MLI Blkts
Therm x30
Data Line
Comm
TxAnt / Radio / TNC
Ext TNC
Serial / I2C
EEDS
Flight Computer
Interface Board
Wireless
Power
Li-Poly Battery
+15V Unreg. Line
+28V Line
Prim SA
Body SA PCB
Mechanisms
EMC Hinges x4
HOP
Tip-Mass Upper LB
Tip-Mass
TM Cam
TMComm
Battery
Tether Release
Mech
TM Lower
LB
PCB
MicroProc
IB
RxAnt / Radio / TNC
Tip-MassAnt / Modem
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ADCS Requirements
Requirement Method Status
The ADCS system must not exceed a mass of 2.25 kg. Design
The ADCS system must operate on less than 4 W. Design
Nadir attitude must be maintained for communication and imaging objectives.
Design, Analysis
Attitude must be determined to within 2 in each axis. Design, Analysis
Attitude must be controlled to within 10 in each axis. Design, Analysis
Torque Rods must lie in right hand orthogonal system. Design
Need an I2C line from C&DH. Design
Detumble the spacecraft within 24 hours of deployment. Design, Analysis
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ADCS Summary• Design
– Sensors• Honeywell HMC2003
Magnetometer• 3 Single-Axis Rate Gyros from
Analog Devices• 12 Single-Axis Ithaco Sun
Sensors– Actuators
• 3 Orthogonal Torque Rods w/ variable current-levels
• Gravity-Gradient Tether– Software
• P-D or LQR Controller• IGRF Magnetic Model• Onboard Orbit Propagator
• Performance– Torque Rods
• m = 3-5 A-m2, P ≈ 2.5 W• Slew rate ≈ 10 min/degree Back to Block
Diagram
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C&DH RequirementsRequirement Method Status
The C&DH system must not exceed a mass of 0.56 kg. Design
The C&DH system must operate on less than 4W. Design
The TNC must have a dedicated RS-232 serial port. Design
The radios must have one RS-232 serial port each. Design
There must be two USB ports for the cameras. Design
Must be able to maintain a 802.11b link between the main satellite and the tip-mass.
Design, Test
C&DH must be able to communicate to EPS through a RS-232 serial port.
Design
There must be enough memory to store all of the software and all of the data.
Design, Test
Provide an interface for all of the thermistors on the spacecraft. Design
An Ethernet port must be available for the GSE. Design
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C&DH Summary
• Design– Motorola PowerPC 823e from
Embedded Planet• SDRAM: 64MB• FLASH: 16MB• NVRAM: 128kB• Real-Time Clock• 3 Serial Ports• 1 USB Port• 1 Ethernet Port• I2C
– Interface Board• Shares USB Port• Shares Serial Ports• Interfaces the thermistors to
the flight computer
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Software
Requirement Method Status
The Software must not use more than 16MB of NVRAM and 16MB of RAM.
Design, Test
The on-board data system shall monitor the health and status of the DINO satellite and generate compressed data summaries.
Design
The flight data system should generate spacecraft health and status data sets for storage and/or downlink.
Design, Analysis
The flight software shall have the capacity to control the operational state of every hardware component like ADCS.
Analysis
Multiple tasks shall be operated concurrently, namely, SCL, FSW tasks, etc.
Analysis
Perform all Mission Operation tasks. Design
The flight software must use Linux as the embedded kernel and operating system.
Design, Test
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Software• PowerPC Architecture• Linux Operating System• C++ and SCL• Reusing 3CS base software• Data Management and Testing
– Concurrent Version System (CVS) for version control, i.e. code identification
– CVS will tag a baseline Release – Testers job to account and audit what goes into a Release – All tests to be completed and satisfactory prior to each Release– CVS to control access to files under CM
Back to Block Diagram
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Communication Requirements
Requirements Method Status
The communication system must not exceed a mass of 1.3 kg. Design
The communication system must operate on less than 32W when receiving and transmitting at the same time.
Design, Test
The communication system must operate on less than 1W while in receive mode only.
Design, Test
Must be capable of two-way communications. Design, Test
A receiving antenna is required for the 2 m band uplink. Design, Test
A transmitting antenna is required for the 70 cm band downlink. Design, Test
Data transfer rates must be high enough to accommodate all data.
Design, Test
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Comm Summary
• Frequencies– Uplink 2m (145 MHz);
Downlink 70cm (436 MHz)
• Design– Two Kenwood TH-D7 radios: dual band 70cm / 2m,
with internal TNCs (1200 baud)– External Timewave PK-96 TNC (9600 baud)– Antenna
• Trade between patch and deployable monopole
• Performance– Data Rate: 9600 baud
• Required Eb/No: 13 dB• Performance: 13.4 dB
– Data Rate: 1200 baud• More reliable and proven
– Two-way communication• Max 50 kB uplink; 25 kB per cloud topo map downlink Back to Block
Diagram
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Thermal RequirementsRequirement Method Status
The thermal system must not exceed a mass of 0.4 kg for the main satellite and 0.1 kg for the tip mass.
Design, Test
The thermal system must not exceed 1 W during lit portion of orbit and 0.2 W during eclipse.
Design
Must keep the Power system between 0 and 40C Design, Analysis
Must keep the C&DH system between 0 and 70C Design, Analysis
Must keep the Comm system between -20 and 60C Design, Analysis
Must keep the ADCS system within a temperature range between -40 and 85C
Design, Analysis
Must keep the Structures system between -60 and 65C Design, Analysis
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Thermal Summary
• Design: Passive– 30 Temperature Sensors– Insulation Possible
• MLI Blankets
– Radiators as needed
• Thermal Analysis– An empty spacecraft
analysis has been complete and indicates a cold spacecraft
– Ball Aerospace assisting with thermal modeling
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Structures Requirements
Requirement Method Status
Entire structures with mechanisms will be less than 9.2 kg Design, Test
Fixed base Natural Frequency must be greater than 100HZ at the Shuttle Interface Plane (SIP).
Design, Test
Center of mass is to be no more than 0.25” from centerline and 12” from the SIP.
Design, Analysis
The completed satellite must fit within ICU envelope. Design
Each component must be less than a length of 8.5” and a height of less than 11”.
Design
Structures must provide a housing for all components. Design
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Structures Summary
• Design– Hexagonal Al iso-grid main
structure• 12.5” tall, 17.76” diameter
• Al 6061
• 5.56 kg
– Component Boxes• 9 boxes
• All mounted on the iso-grid
– Finite Element Analysis is in the first stages
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Mass Budget
Main Satellite Mass Budget
ADCS9%
Solar Panels
8%
Power8%
Comm5%Margin
25%
Str/Mech36%
Thermal2%
C&DH2%
Systems3%
Science2%
Allocation (%) Budget (kg)
Subsystems
ADCS 9 2.25
C&DH 2 0.56
Comm 5 1.13
Power 8 2.06
Science 2 0.56
Software 0 0.00
Solar Panels 8 2.06
Str/Mech 36 9.00
Thermal 2 0.38
Cabling 3 0.75
Total 75% 18.75 kg
Margin 25% 6.25 kg
Total 100% 25.00 kg
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Mass Budget
Tip-Mass Mass Budget
Comm8%
Science9%
Systems2%
Thermal2%
Power15%Str/Mech
39%
Margin25%
Allocation (%) Budget (kg)
Subsystems
ADCS 0 0.00
C&DH 0 0.00
Comm 8 0.38
Power 15 0.75
Science 9 0.45
Software 0 0.00
Solar Panels 0 0.00
Str/Mech 39 1.99
Thermal 2 0.08
Cabling 2 0.11
Total 75% 3.75 kg
Margin 25% 1.25 kg
Total 100% 5.00 kg
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Power Budget
Normal Operating Daytime
Science5%
Comm Tx6%
Comm Rx4%
Thermal3%
ADCS13%
C&DH13%
Power31%
Margin25%
Allocation (%) Budget (W-hr)
Subsystems
ADCS 13 3.9
C&DH 13 3.9
Comm 10 2.8
Power/Power Losses 31 9.2
Science 5 1.5
Software 0 0.0
Solar Panels 0 0.0
Str/Mech 0 0.0
Thermal 3 0.9
Cabling 0 0.0
Total 75% 22 W-hr
Margin 25% 5.5 W-hr
Total 100% 27.5 W-hr
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Power Budget
Tip-Mass
Power17%
Comm45%
Imaging11%
C&DH2%
Margin25%
Allocation (%) Budget (kg)
Subsystems
ADCS 0 0.0
C&DH 2 0.2
Comm 45 4.0
Power 17 1.5
Science 11 1.0
Software 0 0.0
Solar Panels 0 0.0
Str/Mech 0 0.0
Thermal 0 0.0
Cabling 0 0.0
Total 75% 6.7 W-hr
Margin 25% 1.7 W-hr
Total 100% 8.4 W-hr
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Project Milestones
Milestone Start Date End Date
External PDR 8/14/03 8/14/03
Prototyping 8/20/03 1/16/04
Peer Review 12/8/03 12/8/03
Internal CDR 2/4/04 2/4/04
Requirements Freeze 2/11/04 2/11/04
System Review 4/28/04 4/28/04
Electrical Integration 5/24/04 8/20/04
External CDR 8/13/04 8/13/04
Mechanical Integration 4/7/04 11/26/04
Deliver to AFRL 1/5/05 1/5/05
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Risk Assessment
AD
CS
C&
DH
Com
m
Ima
ging
Mechanism
s
Pow
er
Softw
are
Structures
Therm
al
Tip-M
ass
Overall P
rogram
Assessm
ent
Performance
Schedule Cost
Safety
Testing
Personnel
Resources
Overall Subsystem Assessment
= low risk
= medium risk
= high risk
= N/A
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix
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Appendix Index
• Appendix A – Requirements• Appendix B – Systems Charts• Appendix C – Management an
d Outreach • Appendix D – Subsystem Bloc
k Diagrams• Appendix E – Test Plans• Appendix F – ADCS• Appendix G – C&DH
• Appendix H – COMM • Appendix I – EPS • Appendix J – Mechanisms • Appendix K – Software • Appendix L – Structures • Appendix M – Tip Mass
• Appendix N – Thermal
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Deployment and Intelligent Nanosat
Operations
Appendix A
Requirements
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ADCS Requirements• The ADCS subsystem shall provide three-axis attitude determination• The ADCS subsystem shall provide three-axis attitude control
– The ADCS subsystem shall provide a way to slew/detumble the spacecraft to within 45 degrees of nadir-pointing and within TBD deg/s of angular rotation prior to the tether’s deployment
– The tether shall be deployed at a rate such that the libration angle does not increase beyond 45 degrees at any point during its deployment
• The ADCS subsystem shall keep the spacecraft within 5 degrees of nadir-pointing at all times after the tether’s deployment
– The tether shall be long enough, given the tip-mass’ mass, to provide a stable nadir-pointing configuration throughout DINO’s entire mission
– The gravity-gradient torque established by the boom shall be larger than the sum of all other torques experienced by the spacecraft
– The tip-mass shall be massive enough, given the tether’s length, to provide a stable nadir-pointing configuration throughout DINO’s entire mission
– The tether shall be resistant to environmental hazards for a duration of 12 months– The tether material shall be resistant to micrometeoroids for 12 months– The tether material shall be resistant to radiation degradation for 12 months– The tether material shall be resistant to changes in the thermal environment, i.e., shall not
stretch more than specified safe• The tether material shall be capable of withstanding the maximum tension forces
experienced during 12 months, namely, the tension experienced after entering the sunlit side of the orbit at end-of-life Back to
Index
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ADCS Requirements• The ADCS subsystem shall provide yaw-control to within +/- 10 ° degrees of
a designated yaw-angle and TBD deg/s within a designated yaw-angular rate
• The ADCS subsystem shall have a mass less than 2.25 kg• The ADCS subsystem shall operate on less than 4 Watts of power• The ADCS subsystem shall operate on 5V and/or 12V power lines• All ADCS subsystem components shall comply with NASA’s safety
requirements– The ADCS components shall comply with electrical bonding regulations– The ADCS wiring shall be sized according to NASA regulations– The ADCS components shall comply with NASA’s outgassing regulations– The tether material shall outgas within specified limits– The ADCS components shall comply with NASA’s corrosion-resistance
specifications– All ADCS components shall meet the requirements for low-risk fracture
classification– Failure of any ADCS component shall not result in a catastrophic hazard to the
Space Shuttle– All ADCS components shall be composed of acceptable materials per NASA
requirements Back to Index
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C&DH Requirements• The C&DH subsystem shall provide the means for each subsystem to
communicate between one another– The C&DH subsystem shall support the ADCS subsystem
• The C&DH subsystem shall either provide continuous processor feedback to the ADCS subsystem or a microprocessor for the ADCS subsystem
• Given a microprocessor, the C&DH subsystem shall be capable of commanding and receiving data from the ADCS microprocessor
• The microprocessor shall communicate through a TBD port– The C&DH subsystem shall support the Comm subsystem
• There shall be one dedicated RS-232 port available for the receiving Comm radio, used to initialize the radio and for all command reception
• There shall be one RS-232 port available for the transmitting radio, used to initialize the radio and when transmitting at 1200 baud
• There shall be one RS-232 port available for the external TNC to transmit science and engineering data to the ground at 9600 baud
• There shall be one RS-232 port available or one Ethernet port available for the CompactRF Industrial Wireless Modem
Back to Index
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C&DH Requirements• The C&DH subsystem shall provide a beacon for the Comm subsystem to
transmit during DINO’s initial deployment phase and during any period of time that DINO is in safe-mode
• The spacecraft shall have a means of identifying, decoding, processing, and error-checking commands
• The system shall be capable of decoding TBD and provide a bit-error rate of no more than 10-5 error-bits/good-bit
• The spacecraft shall have a means of error-checking and encoding the science and engineering data prior to communication with ground
• The system shall be capable of identifying and repairing errors (?)• The transmissions shall be encoded using a TBD-encoding scheme• The C&DH subsystem shall support the Power subsystem• C&DH will command EPS through a RS-232 serial port.
– The C&DH subsystem shall support the Science subsystem• Commands need to be sent through two USB ports.
– The C&DH subsystem shall support the Structures/Mechanisms subsystem• A TBD number of data lines are needed to determine If the deployments are
fully deployed.– The C&DH subsystem shall support the Thermal subsystem
• The C&DH subsystem shall either provide continuous processor feedback to the Thermal subsystem or a microprocessor for the Thermal subsystem
Back to Index
Colorado Space Grant Consortium 57
DINO PDR April 21, 2023
C&DH Requirements• The C&DH subsystem shall provide data storage capability for all science
and engineering data between ground communication opportunities– The C&DH subsystem shall provide for on-board storage of flight data
• The C&DH subsystem shall provide the necessary operating system, memory, and storage space to operate the flight software
– The flight data system shall include the Linux operating systems– The flight data system shall include a flight computer– The flight computer should be tolerant of radiation (budget pending)– The flight computer shall be able to be reset
• The C&DH subsystem shall be less than 0.56 kg in mass• The C&DH subsystem shall operate with less than 4.0 Watts of power• The C&DH subsystem shall operate on 5V and/or 12V lines• The C&DH subsystem shall support the components on the tip-mass
– The tip-mass communication system shall have a means to check all communications for errors
– The primary spacecraft shall have error-checking software– The tip-mass shall re-transmit all commands to the primary spacecraft to check
for errorsBack to Index
Colorado Space Grant Consortium 58
DINO PDR April 21, 2023
COMM Requirements• The Comm subsystem shall be capable of receiving commands from
ground at any time during its mission– The Comm subsystem shall provide a receiving antenna and radio dedicated
for communication with the ground• The primary receiving system shall be powered at all times
– The data system should support an uplink data rate of at least TBD– The communication system shall be sensitive to frequencies between TBD and
TBD (145 +/- 1 MHz)• The Comm subsystem shall be capable of transmitting all science and
engineering data to the ground– The Comm subsystem shall provide a transmitting antenna and radio dedicated
for communication with the ground– The Comm subsystem shall provide a sufficiently large transmission rate to
transmit all pertinent science and engineering data to the ground using only the available communication opportunities
– The data system should support a downlink data rate of at least TBD– The transmissions shall be contained in the (TBD) band (436 +/- 1 MHz)
Back to Index
Colorado Space Grant Consortium 59
DINO PDR April 21, 2023
COMM Requirements• The Comm subsystem shall have the capability of communicating with the ground at a minimal
level in any attitude configuration• The Comm subsystem shall be capable of two-way communication with the tip-mass
experiments– The tip-mass shall have at least one antenna capable of receiving commands and transmitting data to
the primary satellite– The spacecraft shall have at least one antenna capable of receiving data and transmitting commands to
the tip-mass experiments– The communication system shall transmit data at a rate of at least TBD kbps with a margin of at least
TBD dB– The communication system shall have a means to check all communications for errors– The primary spacecraft shall have error-checking software– The tip-mass shall re-transmit all commands to the primary spacecraft to check for errors– The wireless communication shall comply with all communication regulations – The wireless communication shall not interfere with other satellite communications– The wireless communication shall comply with all ground communication regulations provided that the
signals can be received on the ground– The Comm subsystem shall provide a CompactRF OEM Industrial Wireless Modem for two-way
communication with the tip-mass experiments• The Comm subsystem shall provide at least one ground communication station• The Comm subsystem shall be less than 1.13kg mass• The Comm subsystem shall operate with less than 25 Watts of power while transmitting• The Comm subsystem shall operate on 5V and/or 12V lines• All Comm subsystem components shall comply with NASA’s safety requirements Back to
Index
Colorado Space Grant Consortium 60
DINO PDR April 21, 2023
EPS Requirements• The power subsystem shall meet all of the spacecraft components’ power needs
– A power system shall be employed consisting of solar cells, batteries, converters, and buses to monitor, regulate, and distribute power to all spacecraft components• The spacecraft shall have solar cells pointed toward the sun at all times when not in the Earth’s shadow.
– MicroSat’s 8%-efficient solar cells shall be implemented in two symmetric FITS solar array systems– Higher-efficient solar cells shall be fixed to the body of the spacecraft, including the top face, the two leading faces, and the
two aerofins– Solar cells not pointing toward the sun shall not drain power from the system– Solar cells shall meet all safety requirements– Wiring harnesses shall be provided to connect DINO to each populated face
• The spacecraft shall have a battery system to maintain power throughout eclipse periods and safe-mode periods– The battery system shall be designed to operate over 12 months, equivalent to about 5,850 orbits– The battery system shall be capable of being fully charged with only the allocated power from the solar cells– The battery system shall comply with all Shuttle safety requirements
• The spacecraft shall have buses and converters to supply regulated power to any and all of its components– The power system shall have the capability to monitor the current and voltages across its buses
• There shall be appropriate power lines connecting each component that requires power to the power system– Each wire shall comply with NASA’s wire-sizing regulations
• The spacecraft shall have the capability to dump excess power from its system • The spacecraft shall have a safe-mode power system if the available power drops below a given threshold.• All wiring and inhibits shall be derated• The power subsystem shall support the ADCS subsystem
– The power subsystem shall support the C&DH subsystem• The MPC 823e flight computer shall receive either 5 V or 3.3 V DC power and at most 1 Amp of current per C&DH hardware requirements
– The power subsystem shall support the Comm subsystem• The transmitter shall receive 1.4 A @ 12 V during transmissions and 90 mA @ 12 V when idle• The receiver shall receive 90 mA @ 5 V continuously• The TNC shall receive 400 mA @ 12 V continuously (no less)• The CompactRF OEM Industrial Wireless Modem shall receive 600 mA @ 4.4 to 5.5 Volts
– The power subsystem shall support the Science subsystem– The power subsystem shall support the Structures/Mechanisms subsystem
• The HOP/spring deployment of the tip-mass/tether system shall receive 10 Watts of power, preferably on a 24V or 28V line• Each of the two HOP mechanisms holding the FITS solar arrays shall receive 10 Watts of power upon deployment, preferably on a 24V or 28V line• Each of the EMC hinges shall receive 10 Watts of power for one minute upon deployment, either on a 24V line or, after redesigning the hinges, on a
12V or greater line Back to Index
Colorado Space Grant Consortium 61
DINO PDR April 21, 2023
EPS Requirements• The power subsystem shall meet all of the tip-mass components’ power needs
– Power to the tip-mass experiments shall be activated immediately prior to tether-deployment
– The tip-mass shall include sufficient battery power to provide power to all tip-mass components for the deployment-phase’s duration
– The tip-mass batteries shall be capable of being charged prior to tether-deployment• The power subsystem shall implement MicroSat’s FITS solar panels• The spacecraft shall meet the electrical design requirements listed in 6.5 of the ICU
User’s Guide• The Power subsystem shall be less than 2.06 kg in mass on the main satellite and
less than 0.75 kg on the tip-mass• The Power subsystem shall operate with less than TBD Watts of power• The Power subsystem shall operate on 5V, +/-12V, and 28V lines• All Power subsystem components shall comply with NASA’s safety requirements
– The tip-mass power system shall comply with NASA’s safety regulations– The tip-mass battery system shall comply with NASA’s safety regulations
Back to Index
Colorado Space Grant Consortium 62
DINO PDR April 21, 2023
Science Requirements• Cloud heights shall be measured using a stereoscopic imaging technique
– Cameras shall be used to image the clouds in the visible spectrum– The cloud-height measurements shall have a resolution of at least 500 meters– The cameras shall have a 40-55 degree field of view– The cameras shall have a shutter speed of 1/60th of a second or faster– Ample amount of time shall be allotted for the software system to finish
processing an image before another image is required– Each image shall contain at least one identifiable cloud feature and at least one
identifiable ground feature– Each of the multiple images used to produce a topographic map of the cloud
features must contain the same cloud features and the same ground features• The spacecraft’s deployments shall be imaged by a camera in the tip-
mass– The tip-mass camera must be oriented such that it can view all deployments– The tether shall not block the field of view of the spacecraft– The tip-mass camera must have sufficient resolution to observe the
deployments of DINO’s structures– The camera shall be able to focus on the spacecraft– The images shall display ample light levels and contrast to see the spacecraft
and its deployables Back to Index
Colorado Space Grant Consortium 63
DINO PDR April 21, 2023
Science Requirements
• The Science subsystem shall be less than 0.56 kg on the main satellite and 0.45 kg on the tip-mass
• The Science subsystem shall operate with less than 11 Watts of power on the main satellite and 2.5 Watts of power on the tip-mass
• The Science subsystem shall operate on 5V and/or 12V lines• All Science subsystem components shall comply with NASA’s
safety requirements – There shall be no pressurized vessels in the science subsystem,
including the lens of each camera– All components shall comply with NASA’s outgassing specifications– Any glass components shall comply with NASA’s regulations– All components shall either be contained or meet NASA’s
requirements to be a low-released mass partBack to Index
Colorado Space Grant Consortium 64
DINO PDR April 21, 2023
Software Requirements• The software team shall provide flight software to operate each subsystem
– The flight software shall have the capacity to control the operational state of every hardware component• The flight software shall be capable of controlling the ADCS subsystem on either the flight computer
or the ADCS microprocessor per ADCS/C&DH requirements• The flight software shall be capable of providing a means to communicate between every
subsystem via the C&DH subsystem• The flight software shall be capable of operating the Comm subsystem
– The flight software shall initialize the receiver radio (frequency, squelch), the transmitter radio (frequency, power output), and the TNC
– The flight software shall be capable of encoding and constructing data packets to be transmitted to the ground– The flight software shall be capable of receiving transmissions from the ground, recognizing the transmissions,
checking those transmissions for errors, decoding the transmissions into their corresponding commands, and processing the commands accordingly
– The flight software shall be capable of acknowledging the successful receipt and/or the successful transmission of communications with the ground
• The flight software shall be capable of supporting the Power subsystem• The flight software shall be capable of supporting the Science subsystem
– The flight computer shall be capable of building a topographic map out of two or more images of the same cloud features
• The flight software shall be capable of controlling the thermal system either on the flight computer or on the thermal microprocessor per Thermal/C&DH requirements
• Multiple operations shall be operated concurrently, namely, CASPER, SCL, FSW, etc…
• All software shall be contained in the following hardware:– either 16 or 64 MB of SDRAM, either 8 or 16 MB of flash memory, and either 0 or 16 MB
of NVRAM per C&DH hardware limitations
Back to Index
Colorado Space Grant Consortium 65
DINO PDR April 21, 2023
Software Requirements• The flight software shall be capable of the following requirements
– The on-board data system shall constraint check commands using SCL– The on-board data system shall monitor the health and status of the DINO satellite– The on-board data system shall provide for scheduled and interactive (immediate) control of spacecraft
operations, payload operations, resource usage, and data transport– On-board operational constraints, rules, sequences, and algorithms shall be updated by users– On-board operational constraints, rules, sequences, target synchronization, and algorithms should be
updated automatically– The on-board data system should support the migration of autonomy through tailoring of flight
applications software and updates to procedures, rules, constraints, etc.– DINO may be able to obtain stereo measurements by determining the observation timing and instrument
pointing– The flight data system should be able to generate compressed data summaries– The fight data system should be able to generate spacecraft health and status data sets for storage
and/or downlink– The on-board operations timeline should be capable of being updated
• The flight software shall be designed to the following requirements– The flight software should be a basic system that can be tailored as the mission evolves– The flight software should include reusable libraries, classes, and structures– The flight software should support processing software for simple target recognition, data compression,
and stereo targeting– The flight software design should support modular software– The flight software should try to make use of previously-existing code– The flight software should try to make use of Linux as the embedded kernel and operating system, SCL
for command, control, and performance evaluation and fault management.
Back to Index
Colorado Space Grant Consortium 66
DINO PDR April 21, 2023
Software Requirements• The software team shall provide software for DINO’s ground support
– The ground data system shall receive orbit ephemeredes– The ground data system shall provide for constraint checking of commands and verification
of command execution– The ground data system shall monitor health and status of the DINO satellite and end-to-end
system (EEDS)– The ground data system shall plan and schedule resource use, spacecraft operations,
imaging operations, and command/data transport activities– Users of the ground data system shall be able to update operational constraints, rules,
sequences, displays, and algorithms– The ground system should be able to autonomously update software constraints, rules,
sequences, target synchronization, displays, and algorithms– The ground data system should provide for the migration of autonomy from manual to
supervised, to automated, and from ground to on-broad through tailoring of applications software and updates to procedures, rules, and scripts
– One user ground data system should be able to control operations of the DINO satellite– One user ground data system should be able to monitor, evaluate performance of, and
request operations for one or more flight subsystems– The ground data system should be able to predict target pointing coordinates and timing; to
synchronize with the satellite overpass; and to select data for downlink– The ground data system shall/provide a flexible interface to the database for various
applications– Access to the DINO mission database should be available through www– Provide ability to query the DINO database using number of search criteria Back to
Index
Colorado Space Grant Consortium 67
DINO PDR April 21, 2023
Software Requirements
• The ground software shall be designed to the following requirements– The ground software should support a UNIX/Linux operating system– The ground software should include reusable libraries, classes, and structure– The ground data systems shall plan and schedule resource use, spacecraft
operations, imaging operations, and command/data transport activities– Users of the ground data system shall be able to update operational
constraints, rules, sequences, displays, and algorithms– The ground data system should provide for the migration of autonomy from
manual to supervised, to automated, and from ground to onboard through tailoring of applications, software, and updates to procedures, rules, and scripts
– The ground data system should be able to predict target pointing coordinates and timing; to synchronize with the satellite overpass; and to select data for downlink
– The ground data system shall provide a flexible interface to the database for various applications
– Access to the DINO mission database should be available through www– The ground data system shall provide the ability to query the DINO database
using a number of search criteria
Back to Index
Colorado Space Grant Consortium 68
DINO PDR April 21, 2023
Structures and Mechanisms Requirements
• The Structure/Mechanism subsystem shall provide the primary structural support for DINO and its subsystems
– DINO’s structure shall comply with all of the structural requirements given in the ICU User’s Guide– The entire spacecraft shall fit inside of the specified ICU physical envelope
• The allowable static envelope is defined as a cylindrical right prism with a diameter of 18.7” (47.5 cm) and a height of 18.7” (47.5 cm)
– The spacecraft shall have a maximum mass of 30 kg, including the bolts used to attach the spacecraft to the NSS at the SIP
– The spacecraft’s center of gravity (CG) shall be within 0.25 inches of the geometric central axis of the ICU.
– The spacecraft’s CG shall not lie more than 12 inches above the satellite interface plane (SIP)– The spacecraft shall have a mechanical interface with the ICU per 6.2.1 and 6.2.2 in the ICU User’s
Guide– The spacecraft shall have a fundamental frequency above 100 Hz given a fixed-base condition at the SIP– The spacecraft shall use materials with high resistance to stress corrosion cracking wherever possible
per 6.3.2 in the ICU User’s Guide– The spacecraft shall be capable of handling all loads given in 6.3.3 of the ICU User’s Guide– DINO’s structure shall reduce the risks of any environmental hazard to acceptable levels– The ADCS subsystem shall have structural support
• The magnetic torquers shall be structurally secure to the main structure– The C&DH subsystem shall have structural support
• The MPC 823e flight computer shall be subjected to less than 2 g’s RMS of vibrations, within 20 and 2000 Hz in frequency at all times
• The MPC 823e flight computer shall be contained within a space with dimensions of 9.5 cm x 11 cm x 2.5 cm or greater• The MPC 823e flight computer shall experience no more than 80% non-condensing humidity at all times during DINO’s
mission Back to Index
Colorado Space Grant Consortium 69
DINO PDR April 21, 2023
Structures and Mechanisms Requirements
– The Comm subsystem shall have structural support• The transmitter is 340 grams including the battery (unknown without the battery)• The transmitter has the dimensions: 12.5 cm x 6.0 cm x 3.0 cm• The receiver is 340 grams including the battery (unknown without the battery)• The receiver has the dimensions: 12.5 cm x 6.0 cm x 3.0 cm• The TNC is 544 grams• The TNC has the dimensions: 15.57 cm x 18.8 cm x 3.43 cm• The two antennas shall be 50 cm and 17.5 cm in the vertical direction
– The Power subsystem shall have structural support– The Science subsystem shall have structural support– The Thermal subsystem shall have structural support
• DINO’s outward-facing sides shall be painted in such a way to provide an improved internal thermal environment
• DINO shall have an outward-facing side dedicated to radiating excess heat into space– The principal inertial axes of the spacecraft shall lie as close as possible to the body-axes
per ADCS requirements• The body-axes are defined in the following way:
1-axis along the velocity-vector of the spacecraft2-axis along the cross-track vector of the spacecraft3-axis along the radial (tether) vector
– The radial axis (along the boom) shall be the minor axisThe axis parallel to the FITS deployments shall be the intermediate axis
– The spacecraft shall be symmetric wherever possible per ADCS requirements– The center of mass (CG) shall lie along the radial axis
Back to Index
Colorado Space Grant Consortium 70
DINO PDR April 21, 2023
Structures and Mechanisms Requirements
• The Structure/Mechanism subsystem shall implement a gravity-gradient boom– The tether system shall not bear any loads prior to its deployment– The tip-mass shall be stowed in a fail-safe fashion– The tip-mass shall be held during launch using Planetary System’s Lightband (TBD)– The tip-mass/tether system shall be the first deployed structure– The Structure/Mechanism subsystem shall ensure a correct deployment of the gravity-
gradient boom, producing a nadir-pointing spacecraft configuration per ADCS requirement– The boom shall be deployed in an open-loop fashion, i.e., spring-loaded with a slack
tether– The tether shall be deployed at a rate such that the libration angle does not increase
beyond 45 degrees at any point during its deployment– The system shall have a damping mechanism to reduce the gravity-gradient oscillations
induced by the final orientation of the tether’s deployment– The tether shall be securely attached to the spacecraft– The tether shall be securely attached to the tip-mass– The boom shall be attached to each mass along each mass’ center of gravity– The tip-off rate of the tip-mass’ deployment shall not result in tether-wrapping
Back to Index
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DINO PDR April 21, 2023
Structures and Mechanisms Requirements• The Structure/Mechanism subsystem shall implement FITS solar arrays
– The FITS solar arrays shall not bear any loads during launch– The FITS solar arrays shall be securely attached to the spacecraft before and after their deployment– All latching mechanisms shall be released prior to deployment– Power shall be provided to release any latching mechanism that requires power– Both solar arrays shall be held during launch using a Frangibolt– The FITS solar arrays shall be deployed after the spacecraft is in a stable nadir-pointing configuration
• The Structure/Mechanism subsystem shall implement elastic memory-composite hinges to deploy the aerofin panels
– There shall be two EMC hinges per aerofin panel deployment– The aerofins shall be securely attached to the main spacecraft before and after deployment– All latching mechanisms shall be released prior to deployment– Power shall be provided to release any latching mechanism that requires power– Each panel shall be held by at least one HOPS mechanism (TBD)– The aerofins shall have the capacity to be populated by body-mounted solar cells– The deployment shall allow for wiring harnesses to run across the panel/satellite interface– The wiring harnesses shall not apply any forces or pressure on the solar cells at any time– The EMC hinges shall provide ample torque to release the aerofin panels including wire harnesses that
span the body-aerofin interfaces– Any solar cells integrated with the aerofin panels shall not be damaged during construction, integration,
transportation, launch, and deployment– All latching mechanisms on the spacecraft shall not press any set of body-mounted solar cells against any
other surface– Spacers shall be used, if necessary, to prevent a populated surface from coming into contact with any other
surface on the spacecraft– Each aerofin system shall be resistant to radiation damage for 6 months Back to
Index
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DINO PDR April 21, 2023
Structures and Mechanisms Requirements• The Structures subsystem shall support the tip-mass components
– All of the tip-mass components shall be contained in a box for structural support and to help protect the components from environmental hazards
– The tip-mass experiments shall be protected from all hazardous radiation for their duration
– The tip-mass experiments shall be protected from all micrometeoroid hazards for their duration
• The Structures/Mechanisms subsystem shall be less than 9 kg for the main satellite and 2 kg in mass for the tip-mass (TBD)
• The Structures/Mechanisms subsystem shall operate with less than TBD Watts of power
• The Structures/Mechanisms subsystem shall operate on 5V, 12V and/or 28V lines
• All Structural/Mechanical subsystem components shall comply with NASA’s safety requirements
– The tether material shall comply with NASA’s outgassing regulations– The FITS solar arrays shall be attached in such a way that they comply with all of
NASA’s structural requirements– The aerofin panels shall be attached in such a way that they comply with all of
NASA’s structural requirements
Back to Index
Colorado Space Grant Consortium 73
DINO PDR April 21, 2023
Thermal Requirements
• The Thermal subsystem shall determine the temperature of all components of the spacecraft during all phases of DINO’s mission
– The Thermal subsystem shall provide Thermisters at every location where the spacecraft’s temperature needs to be determined
• The Thermal subsystem shall keep each component within its specified temperature range, be it an operational or non-operational mode3.1.2.1, 3.3.3.1
– The Thermal subsystem shall support the ADCS subsystem– The Thermal subsystem shall support the C&DH subsystem
• The MPC 823e flight computer shall be kept within the operational temperature range of 0°C to 70°C
– The Thermal subsystem shall support the Comm subsystem• The transmitter shall be kept between -20°C and 60°C• The receiver shall be kept between -20°C and 60°C• The TNC shall be kept between -20°C and 60°C
– The Thermal subsystem shall support the Power subsystem• The batteries shall be kept between 0°C and 40°C.
– The Thermal subsystem shall support the Science subsystem– The Thermal subsystem shall support the Structures/Mechanisms subsystem Back to
Index
Colorado Space Grant Consortium 74
DINO PDR April 21, 2023
Thermal Requirements
• The spacecraft shall meet the thermal design requirements listed in 6.4 of the ICU User’s Guide3.2.1.4
• The Thermal subsystem shall be less than 0.4 kg of mass on the main satellite and 0.1 kg of mass on the tip-mass
• The Thermal subsystem shall operate with less than 1.0 Watts of power during the lit portions of the orbits and 0.2 Watts during eclipse
• The Thermal subsystem shall operate on 5V and/or 12V lines
• All Thermal subsystem components shall comply with NASA’s safety requirements
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix B
Systems Charts
Colorado Space Grant Consortium 76
DINO PDR April 21, 2023
Deployable Power Profile
0 50 100 150 200 250 3000
2
4
6
8
10
12
14Total Power Required During Detumble Mode, no comm
time (min)
Pow
er (
W)
Power RequiredAvailable SA PowerOn Battery
0 50 100 150 200 250 3000
10
20
30
40
50
60
70
Battery Capacity During Detumble Mode, no comm
time (min)
Cap
acity
(W
-hr)
Battery CapacityOn Battery
Detumble ModeBack to Index
Colorado Space Grant Consortium 77
DINO PDR April 21, 2023
Deployable Power Profile
Tip-Mass Deployment Mode
0 50 100 150 200 250 300 350 4000
5
10
15
20
25Total Power Required During Tip-Mass Deployment Mode
time (min)
Pow
er (
W)
Power RequiredAvailable SA PowerOn Battery
0 50 100 150 200 250 3000
10
20
30
40
50
60
70
Battery Capacity During Tip-Mass Deployment Mode
time (min)
Cap
acity
(W
-hr)
Battery CapacityOn Battery
Back to Index
Colorado Space Grant Consortium 78
DINO PDR April 21, 2023
Deployable Power Profile
Quick Deployment Case
0 100 200 300 400 500 600 7000
10
20
30
40
50
60
70Total Power Required During Deployments
time (min)
Pow
er (
W)
Power RequiredAvailable SA PowerOn Battery
0 100 200 300 400 500 600 7000
10
20
30
40
50
60
70
Battery Capacity During Deployments
time (min)
Cap
acity
(W
-hr)
Battery CapacityOn Battery
Back to Index
Colorado Space Grant Consortium 79
DINO PDR April 21, 2023
DINO Power Consumption During a Regular Cycle
0
5
10
15
20
25
30
35
40
0 10 20 30 40 50 60 70 80 90 100
Time, Minutes
Pow
er C
onsu
med
, Wat
ts ADCS
C&DH
COMM
Science
Power
Thermal
Total
NOTE: Plot begins and ends as DINO exits eclipse
Eclipse Begins
Back to Index
Colorado Space Grant Consortium 80
DINO PDR April 21, 2023
DINO Available Power Over 1 Year
0 20 40 60 80 100 120 140 160 180 2000
10
20
30
40
50
60
70
80Total Wattage Produced by Solar Panels for 4 Sun Positions in ISS Orbit
time (min)
Pow
er
Outp
ut
(W)
Winter: 23.60 WSpring: 27.28 WSummer: 23.64 WFall: 27.21 W
Nominal max.Power Consumed
Total Wattage Produced by Solar Panels for 4 Sun Positions in ISS Orbit
Pow
er O
utpu
t, W
atts
Time, MinutesBack to Index
Colorado Space Grant Consortium 81
DINO PDR April 21, 2023
Electrical Power System Block Diagram
EPS
Control Card
C&DH
Battery
Prim SA
Back SA
GSE
RS-232 I/O
EMC 2
EMC 1
Boom
EMC 3
EMC 4
HOP
Mag.
TR 3
TR 2
TR 1
Sun Sen.
Radio 1
Radio 2
TNC
HOP SW
Wireless
Cam 2
Cam 1
Imaging
COMM
ADCS
STRUCTURES
SCIENCE
15V
28V
12V
5V5V
Back to Index
Colorado Space Grant Consortium 82
DINO PDR April 21, 2023
Power Budget
Normal Nighttime
Comm Tx
10%
ADCS11%
Comm Rx6%
Thermal1%
Power22%
C&DH25%
Margin25%
Allocation (%) Budget (W-hr)
Subsystems
ADCS 11 1.0
C&DH 25 2.3
Comm 16 1.5
Power/Power Losses 22 2.0
Science 0 0.0
Software 0 0.0
Solar Panels 0 0.0
Str/Mech 0 0.0
Thermal 1 0.1
Cabling 0 0.0
Total 75% 7 W-hr
Margin 25% 1.8 W-hr
Total 100% 8.8 W-hr Back to Index
Colorado Space Grant Consortium 83
DINO PDR April 21, 2023
Power Budget
Safe Mode
C&DH31%
Power34%
Comm Rx9%
Thermal1%
Margin25%
Allocation (%) Budget (W-hr)
Subsystems
ADCS 0 0.0
C&DH 31 3.3
Comm 9 1.0
Power/Power Losses 34 3.6
Science 0 0.0
Software 0 0.0
Solar Panels 0 0.0
Str/Mech 0 0.0
Thermal 1 0.1
Cabling 0 0.0
Total 75% 8 W-hr
Margin 25% 2 W-hr
Total 100% 10 W-hrBack to Index
Colorado Space Grant Consortium 84
DINO PDR April 21, 2023
Overall Mission Timeline
Laun
ch
ICU
Dep
loym
ent
DIN
O D
eplo
ymen
t
Mis
sion
Act
ivat
ion
Tet
her
Dep
loym
ent
FIT
S D
eplo
ymen
t
Aer
ofin
Dep
loym
ent
Nor
mal
Sci
ence
Ope
ratio
ns
Back to Index
Colorado Space Grant Consortium 85
DINO PDR April 21, 2023
Launch Phase
Sys
tem
Che
ck
Laun
ch
ICU
Dep
loym
ent
from
Orb
iter
DIN
O
Dep
loym
ent
Mis
sion
A
ctiv
atio
n
ICU
Sep
arat
ion
Back to Index
Colorado Space Grant Consortium 86
DINO PDR April 21, 2023
Mission Activation Phase
Mis
sio
n A
ctiv
atio
n
Po
wer
Up
Orbit
SunlitEclipse
Flig
ht S
yste
m P
owe
r up
Flig
ht C
ompu
ter
Pow
er U
p
Po
we
r S
yste
m A
ctiv
ate
d
Com
m,
Th
erm
al,
AD
CS
Po
we
r U
p
Sys
tem
Hea
lth C
heck
AD
CS
Act
ivat
ion
Co
mm
un
icat
ion C
omm
In
itia
lizat
ion
Bea
con
Tra
nsm
issi
on
Gro
und
Com
mun
icat
ion
Dep
loym
ent
Ph
ase
Back to Index
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DINO PDR April 21, 2023
Deployment Phase
Tet
her
Dep
loym
ent
Orbit
SunlitEclipse
Tet
her
Sys
tem
Hea
lth C
heck
Att
itude
Sta
biliz
atio
n C
hec
k
Tet
her
Dep
loym
ent
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Up
Tip
-Mas
s H
ealth
Che
ck
Tet
her
Sys
tem
Hea
lth C
heck
Tet
her
Im
age
Acq
uisi
tion
and
D
own
link
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Do
wn
AD
CS
Sta
bili
zati
on
Back to Index
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DINO PDR April 21, 2023
Deployment Phase
FIT
S D
eplo
ymen
t
Orbit
SunlitEclipse
Att
itude
Sta
biliz
atio
n C
hec
k
So
lar
Arr
ay A
ctiv
atio
n
FIT
S D
epl
oym
ent
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Up
Tip
-Mas
s H
ealth
Che
ck
Flig
ht S
yste
m H
ealth
Che
ck
Tip
-Mas
s Im
age
Dow
nlin
k
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Do
wn
Tip
-Mas
s Im
age
Acq
uisi
tion
Back to Index
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DINO PDR April 21, 2023
Deployment Phase
Aer
ofi
n D
eplo
ymen
t
Orbit
SunlitEclipse
Att
itude
Sta
biliz
atio
n C
hec
k
EM
C H
inge
Dep
loym
ent
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Up
Tip
-Mas
s H
ealth
Che
ck
Flig
ht S
yste
m H
ealth
Che
ck
Tip
-Mas
s Im
age
Dow
nlin
k
Gro
und
Com
mun
icat
ion
Tip
-Mas
s P
ower
Do
wn
Tip
-Mas
s Im
age
Acq
uisi
tion
Back to Index
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DINO PDR April 21, 2023
Science Operations Phase
No
rmal
Op
erat
ion
s
Orbit
SunlitEclipse
Att
itude
Sta
biliz
atio
n
Sci
ence
Im
age
Acq
uisi
tion
Gro
und
Com
mun
icat
ion
Flig
ht S
yste
m H
ealth
Che
ck
Bat
tery
Cha
rgin
g
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix CManagement and
Outreach
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DINO Outreach• The Space Grant Outreach Team will develop K-12
modules based on concepts from the DINO mission.• The first module ready for prototyping will be done in
early Fall 2003.• A 5th grade DINO module will be prototyped at Spangler
Elementary in Longmont, Colorado with students in the Mathematics, Engineering, and Science Achievement (MESA) Program in Fall 2003.
• Modules will continue to be created and prototyped with MESA students throughout the remainder to 2003 and 2004.
• The goal is for a complete K-6 curriculum by early 2005.Back to Index
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DINO Outreach
• What is a Module?
A module is an inquiry based, hands-on activity designed to enhance teacher’s lesson plans and student’s comprehension.
• Focus of DINO modules:
1) Satellites as a medium for communication, exploration, and research.
2) Engineering as an important profession that requires team work, critical thinking skills, and problem solving capabilities.
3) Space science in relation to DINO. Back to Index
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Schedule
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Organization Chart
ADCSSteve Stankevich (L)
Jeff Parker
PowerMike Wong (L)
Kevin McWilliams
COMMZach Allen (L)Hosam Gusim
Project Management and Control
Jen Michels (L)Jeff Parker
StructuresTim Shilling (L)Anthony Lowrey
Jen GetzTerry Song
ThermalJosh Stamps (L)Robin Hegedus
Ball (I)
C&DHMike Li (L)
Yosef Alyosef
MOPSSteve Stankevich (L)
Science Jessica Pipis (L)Anders Fornberg
Mike Wong
Systems EngineeringJeff Parker (L)
Anthony Lowrey
SoftwareCory MaccarroneCameron Hatcher
Nick Pulaski
StaffDave BeckwithElaine HansenChris Koehler
Steve Wichman
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Contingency PlansRisk Level Responsible Party Contingency Plan
FITS delivery failure High Systems Team Increase efficiency of body-mounted cells; double-deployment of aerofins
EMC hinge / aero-fin delivery Med Systems Team Eliminate aerofins
Body-mounted solar panel delivery
Med Power Purchase through / provided by other vendor
Power system delivery Med Power Implement in-house design of power system
Tether deployment failure Med Str/Mech Use magnetic torque rods as primary ADCS control
FITS deployment failure Med Str/Mech Go into low-power mission mode
Tether stability, orientation failure
High ADCS Go into failed-stability mission mode; cut tether
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix D
Subsystem Block Diagrams
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MagnetometerHoneywell HMC2003
12V@20mA w/ 40μG resolution
Flight Computer
Power
Att. Det.
IGRF Magnetic ModelOrbit Propagator
Compare ExpectedAnd actualB Fields
Damp rates
Controller
Likely a P-D or LQR
Output cmds to turn rods on/off and current
direction
Possibly use multiple voltage levels requiring a
D/A Converter
Rate Gyro(s)3 single-axis gyros+5V input @ 6mA
Sun SensorsPossible donation by Ithaco
12 single axis
ADCS ElectronicsX Analog
InputsA/D
Conversions
D/A ConversionsTorquer Analog
Outputs
ADCS software running at 1 – 10 Hz
Torque Rods(3) 3/4’’ x 10’’ @ max 150mA nominal
I2C data(3) 0-5V analog
0-300mA
Standard Commands
(3) 0-5V analog
Hardware Flow Diagram12V and 5V supply to board
I2C or I/O lines
Back to Index
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USB RESET SMC1 SMC2 I2C PCMCIA
General Purpose I/O Pins
FLIGHT COMPUTER (5V, 1A)
CPLDCPLD
(5V, 150mA)
USB
Interface
SCIENCERS232 DRIVER
(5V, 7mA)
MULTIPLEXER
(5V)
RADIO TNC
COMM
EPS ADCS
Multiple Wires
Wireless
INTERFACE BOARD
TIP-MASS
RS232 Serial
802.11b(5V, 1A)
System Block Diagram
THERMALINTERFACE
THERMAL
Back to Index
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DINO PDR April 21, 2023
Subsystem Block Diagram
• All Data lines use RS-232 Serial
• Transmitter operates at 12 V when active, 5 V idle.
• Receiver operates at 5 V.• External TNC: 12 V line,
allows 9,600 bps link to ground.
• Internal TNC proven to be reliable at 1,200 bps.
12 V line voltage(5 V when idle)1.4 A16.8 W transmitting(450 mW idle)
5 V line voltage90 mA450 mW(constant)
12 V line voltage400 mA, 4.8 W(constant)
RS-232 Serial9,600 bps(during setup only)
RS-232 Serial9,600 bps(during setup only)
RS-232 Serial9,600 bps(constant)
Transmitter
Kenwood TH-D7
Receiver
Kenwood TH-D7
External Terminal Node Controller(TNC)
InternalTNC
InternalTNC
Antenna
Power LineData Line
Legend
Back to Index
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Electrical Power System Control Card Block Diagram
• Provides SA switch interface• Provide charge control/Monitor• Provides C&DH RS232 interface• Provides secondary Power
• 5V• +/-12V• 28V
• Provides 16 load switches to the spacecraft
SA
Interface
FITS
BodyMount
Charge Control &Monitoring
PMAD
5V Reg
+/- 12VReg
28V Reg
Subsystem
Interface
Battery
C&DH (RS-232)Cmd/Monitor line
Power line
EPS Control Card
Back to Index
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Flow Charts – FSW Uplink
= H/W Interface
Ro
ute
_cm
dS
CL
DB
----
---I
OB
LK
----
--
I2C_mgr
serialmgr
usbmgr
pcmcia/802mgr
MCMD
CMDIN
ADCS
SCI
SWM
POWER
SCL RTE
----
---C
MD
BL
K--
----
BPGEN/COMM
Back to Index
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DINO PDR April 21, 2023
MOPS Flow Diagram
Orbit Event Timeline
Dist. Users Data Dist.
Users
Science Product
Generation
ADCS Analysis
Data Process and Storage
Mission Planning
Sequencing
Command and Control
Back to Index
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MOPS Flow Diagram
Sequence Execution
Sequence Adjustment
Command and Control
Data ProcessingFault Handling and Response
S/C Subsystems
commands
Raw DataBack to Index
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DINO PDR April 21, 2023
Flow Charts – FSW Downlink
= H/W Interface
ADCS
BPGEN/COMM
SCI
SWM
POWER
Ret
urn
_rep
ly
TMOUT
SCL RTE
SC
L D
B
----
---C
MD
BL
K--
----
----
---I
OB
LK
----
-- I2C_mgr
serialmgr
usbmgr
pcmcia/802mgr
MCMD
Back to Index
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DINO PDR April 21, 2023
RTE
S/C Model
Flow Charts - SCL
CMDIN TMOUT
database
Scripts
Incoming
data
Outgoing
data
Rules
FSW
Back to Index
Colorado Space Grant Consortium 107
DINO PDR April 21, 2023
Flow Charts – VMOCC Uplink
MySQL
STK
Disk/ Files
SCLRTE
TMOUT
Webbrowser
ScheduledCmds
Events
Schedules
Schedules
Cmds Pkts
= H/W Interface
Immediate CmdsBack to Index
Colorado Space Grant Consortium 108
DINO PDR April 21, 2023
Flow Charts – VMOCC Dowlink
Disk/ Files
MySQL
STK
SCLRTE
CMDIN
Webbrowser
Sensor data
Orbit data
Pictures
Sci data (pictures)
Pkts
H&S (ADCS)
H&S (All)
Sci data
= H/W Interface
Back to Index
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DINO PDR April 21, 2023
General Layout of Tip Mass
5 VCOMM
Main Satellite Flight Computer
Serial Camera
Power
Sep Switch
Imaging-FPGA
RS
-232
802.11b
Trig
ger5 V
On
RS
-232
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix E
Test Plans
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Integration and Test PlanPower
IntegrationComm.
Integration
Installationin Boxes
Verify
Verify
FunctionalTesting
SubsystemTesting
Mountingon Isogrid
Installation ofMechanisms
on Isogrid
RouteWiring
Harness
CloseStructure
FunctionalTesting
MechanismsTesting
Mount SolarPanels
Verify SolarPanels
FunctionalTesting
ElectricalIntegration
MechanicalIntegration Back to I
ndex
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DINO PDR April 21, 2023
ADCS – Test Plan
• Control algorithm may be tested with an external power supply to electronics board, software running on a linux PC, and mock sensor inputs.
• Attitude determination algorithm may also be tested with linux PC and mock sensor inputs.
• Actual output from actuators can be measured and compared to simulations using the same mock sensor inputs.
• Complete testing after s/c integration is more complicated because the torque rods will not rotate the s/c in a gravity environment. Back to I
ndex
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DINO PDR April 21, 2023
C&DH – Test Plans
• ADC testing requires accompanying software. • All other interface board functionality can be
confirmed with logic analyzer and multimeters.• Testing occurs through the Ethernet port.
Back to Index
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COMM – Test Plan
• Set up ground station in flight configuration• Setup spacecraft in flight configuration (deploy
antennas if necessary)• Transfer files to and from spacecraft CDH
system• Measure throughput• Adjust link as necessary
Back to Index
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DINO PDR April 21, 2023
EPS – Test Plan• Battery to be certified by either Ball Aerospace & Technologies Corp. (BATC) or
cell vendor• Circuitry developed by sub-contractor• Thermal-vacuum and vibration testing will be carried out by CSGC students using
BATC facilities• GSE developed by CSGC
• Verify system inhibits• Verify voltage min/max• Provide SA simulation• Verify battery charge monitors• Verify commands/monitors• Perform step load testing• Monitor temperatures• Environmental testing• Interface with monitor• Exercise all inputs,outputs, and logic operations of EPS
Credit: NASA/JPL http://mars.jpl.nasa.gov/mer/
Back to Index
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FITS Release System – Test Plan
Gravity Back to Index
Colorado Space Grant Consortium 117
DINO PDR April 21, 2023
Aerofins – Test Plan
Aerofin Structure
Sled Air table
CTD Hinge
StructureAir table
Sled
Back to Index
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Software – Test Plan
• Test Harness for each component– A suite of tests – All possible input values and test for the right output
• Minimum and maximum inputs
• Day-in-the-Life (i.e. typical) outputs
– Typically a hardware test platform with software support
• Test Plan for each command– A suite of tests – All possible input values and test for the right output
• Minimum and maximum inputs
• Day-in-the-Life (i.e. typical) outputs
– Typically a software test platform with hardware support Back to Index
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Structures – Test Plan
• FEA (Finite Element Analysis)– Based on CAPE requirements– Guidance provided in UN-SPEC-12311, Stress Analysis Guidelines.
• Sine Sweep Test– 20-2000Hz at .25g– Verified through modal surveys and sine sweep vibration– Conducted before and after sine burst and random vibe tests
• Envelope Verification• Mass Properties
Back to Index
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DINO PDR April 21, 2023
Tip Mass – Test Plan
• COMM– Transmit and receive from both FPGA and flight computer
• Structures– Correct center of mass– Most testing will be taken from main satellite
• Science– Image quality for objects at 20 meter– Proper communications with FPGA
• Power– All system power test to ensure 2 hour operation
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix F
Attitude and Determination Control System
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Magnetometer
Honeywell HMC2003
• 20mA @ 12V
• mass < 100g
• -40 to 85 C operating temp.
• 40 μGauss Resolution
• $200
• 3 Analog Outputs (Bx, By, Bz)
• Set/Reset Capabilities Back to Index
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DINO PDR April 21, 2023
Rate Gyros
Analog Devices ADXRS150
– Single axis rate gyros provide the rotational rate of the s/c about the output axis
– Microchip operating at 5V and 6mA.– Single analog output– -40 to 85°C operating temp– $33 each
Back to Index
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Gravity Gradient
• Gravity gradient provides a restoring torque when a disturbance torque causes a movement from local vertical.
• The torque produced is dependent upon the s/c moments of inertia. This shall be a design concern for placement of s/c components.
• Maximum disturbance torque is Aerodynamic Drag– Daytime during Solar max
– τDrag = 5.56 x 10-5 Nm
• Solar Radiation Pressure– 4.37 x 10-6 Nm
• Magnetic disturbance torques are not considered as disturbances because active magnetic control will be utilized.
Back to Index
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Magnetic Torque Rods
• Ferrite Material wound with wire• Produces a dipole moment that interacts with Earth’s
magnetic field.• Will be designed in-house (unless donated)
Back to Index
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Torque Rod Optimization
3 4 5 6 7 8 9 10 11 12 130.4
0.5
0.6
0.7
0.8
0.9
1
1.1
1.2
Length of the Coil (in)
Dia
met
er o
f the
coi
l (in
)
Circular Torquer-Sizing, I=300 mA, 24 Gauge Wire, =800, #Layers = 10
0.5
1
1
1.5
1.5
2
1 1
2
2 2
3
3
3
5
5 5
7
7
10
10
14
14
19
19
25
0.1 0.2
0.2
0.3
0.3
0.4
0.4
0.4
0.75
20002000
3000
4000 5000
6000
Power Diss (W)Moment (A-m2)Torquer Mass (kg)#Turns
Design• Common ferrite material 33
with μ=800• 24 Gauge Wire• 3-10 Wrappings• 7”-10” long• 1/2” - 3/4” diameter• 0.2 – 0.4 kg each• Max Power: 1.5 W each (300
mA) • 0.75 W Dissipation each• Complete manufacture under
$100 each• After detumbling normal use
should not exceed 150mA• The bigger the better Back to I
ndex
Colorado Space Grant Consortium 127
DINO PDR April 21, 2023
ADC Electronics Board
5V from Power GND 12V from Power
A/D Converter
A/D Converter
I2C dataTo Flight Computer
Mag Sensor
Rate Gyros
Analog
I/O fr
om F
C
Multiplexer
Multiplexer
Multiplexer
ResistorBank
ResistorBank
ResistorBank
To Torque Rods 1,2 and 3
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix G
Command and Data Handling
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Power Consumption
PartPower Consumption
(Watts)
Flight Computer 3
Wireless Interface 3
Miscellaneous Interfaces 1(max)
Total 7(max)
Back to Index
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DINO PDR April 21, 2023
RPX Lite
• SDRAM: 64M• FLASH: 16M• NVRAM: 128K• Real-Time Clock• Interfaces:
– USB
– RS-232 Serial
– PCMCIA
– Ethernet: 10BaseT
– SPI, I2C …
Back to Index
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DINO PDR April 21, 2023
USB Interface
+5V
GND
D+
D-
Camera0
+5V
GND
D+
D-
Camera1
MUX
D+
D-
SelectVcc
Back to Index
Colorado Space Grant Consortium 132
DINO PDR April 21, 2023
Thermal Interface
Sensors in different parts of DINO
Temp. Sensor
Flight Computer
ADC
Sensor select for temperature reading
MUX
Digital Temp. readings
select
Temp. Sensor
+5V
10k
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix H
Communications System
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Illustration of Patch Design
Metal Patch
Dielectric Substrate
r = 37 ± 1
~25 cm
~10 cm
145 MHz coax feed436 MHz coax feed
Mount on Nadir Plate
Back to Index
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Analysis: Power Requirements
• Daytime Operation– Receiver: 0.45 W (5 V, 90 mA) always.– TNC: 4.8 W (12 V, 400 mA) always.– Transmitter: 16.8 W (12 V, 1.4 A) for approx. 2 minutes,
otherwise same as Receiver (0.45 W).
• Nighttime Operation– Receiver: 0.45 W (5 V, 90 mA) always.– TNC: 4.8 W (12 V, 400 mA) always.– Transmitter: 16.8 W (12 V, 1.4 A) for approx. 4 seconds,
otherwise same as Receiver (0.45 W).
• Safe Mode– Same as nighttime. Back to I
ndex
Colorado Space Grant Consortium 136
DINO PDR April 21, 2023
Analysis: Calculating Transmission Time
• We need to find the transmission time in order to find the exact power requirements over the course of one day.
• Time needed to send one packet:– 10 bits/byte * 256 bytes/packet 1200 bits/sec =
2.133 sec/packet• Total transmission time (assuming 25 kB per pass during
daytime):– 2.133 sec/packet * 25 kB/pass
256 bytes/packet = 208.3 sec/pass = ~ 3.5 minutes (absolute minimum)
– May be approx. twice the minimum (resending, errors, etc.)
– This is a realizable amount of time. Back to Index
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DINO PDR April 21, 2023
Analysis: Link BudgetElement Parameter Value Units
Transmitter (Tx)1 Tx Power, Pt 37.0 dBm2 Tx Component Line Losses, Ltl 0.5 dB 3 Tx Antenna Gain (Peak), Gt 2.0 dBi4 Tx Pointing Loss, Ltp 0.0 dB5 Tx Radome Loss, Ltr 0.0 dB
6 EIRP (1-2+3-4-5) 38.5 dBm
PropagationTransmission Frequency, f 436.0 MHzSlant Angle 5.0 degOrbital Altitude, h 425.0 kmLink Range, R 12660.6 kmPropagation Factor, n 1.0
7 Free Space Loss, Ls 167.3 dB8 Atmospheric Absorption, Lpa 0.5 dB9 Precipitation Absorption, Lpp 0.5 dB
10 Total Propagation Loss (7+8+9) 168.3 dB
Receiver (Rx)11 Rx Antenna Gain (Peak), Gr 15.0 dB12 Rx Polarization Loss, Lrpol 3.0 dB13 Rx Pointing Loss, Lrp 0.0 dB14 Rx Radome Loss, Lrr 0.0 dB15 Rx Component Line Losses, Lrl 0.5 dB16 Rx Implementation Losses, Lri 1.0 dB
17 Received effective carrier power -110.8 dBm(6-10+11-12-13-14-15-16)
Noise18 Standard Thermal Noise, kT -174.0 dBm/Hz19 Rx Noise Bandwidth, B 43.0 dBHz20 Rx Noise Figure, NF 10.0 dB
21 Effective Noise Power (18+19+20) -121.0 dBm
Results22 Available CNR (17-21) 10.2 dB23 Data Rate 39.8 dBHz24 Available Eb/No (22+19-23) 13.4 dB25 Implementation Losses 2.0 dB26 Required Eb/No 13.0 dB27 Margin (24-25-26) 6.4 dB
Link Budget Form courtesy of Dr. Stephen Horan, New Mexico State University. Back to I
ndex
Colorado Space Grant Consortium 138
DINO PDR April 21, 2023
Link Budget (cont.)
• Margins for different slant angles– 5 deg: 6.4 dB– 12.5 deg: 9.1 dB– 15 deg: 10.1 dB
Diagram Reference:Vincent L. Pisacane and Robert C.
Moore, Eds., Fundamentals of Space Systems. New York: Oxford University Press, 1994.
Back to Index
DINO PDR April 21, 2023
Deployment and Intelligent Nanosat
Operations
Appendix I
Electrical Power System
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Switch List
FunctionNumber of Switches
Amps Volts
ADCS 3 0.500 5
COMM RX 1 1.500 12
Thermal 3 (TBD) (TBD) (TBD)
Cam-1 1 0.200 5
Cam-2 1 0.200 5
HOPS 1 0.643 28
Memory Hinges 4 0.357 28
Patch Antenna 1 (TBD) (TBD)
COMM TX 1 0.100 5Back to Index
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Deployed driving requirements - Power - 13 Vdc and 60 Watts @EOL
Deployed Solar Array1.10 m2 / Wing Fold Integrated Thin Film Stiffener (FITS) Stainless Steel CIGS Array
85 Watts BOL - AMODeployed Solar Array Meets All Requirements
1.257 m
0.439 m0.439 m
1.636 m
FITS System - DeployedFITS System - Deployed
Back to Index
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Inhibits Layout Diagram
•Main EPS uses separation switches from ICU to trigger system start-up•Connect to two FETs on the high leg of the battery, one on the ground leg•Method to get power into EPS with un-charged battery is via a dedicated SA string
•Solar Array (SA) string switches will act as SA inhibits•Subsystem load switches will act as spacecraft load inhibits
EPS BoardBack to Index
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DINO PDR April 21, 2023
1234
5
67
89
1011
121314
15161718
12345
6
78
9
10
1112
13
14
1516
17
18
19202122
23242526
J1 - FlightTest J2
T
T
T
T
T
T
T
T
Cell-1
Cell-2
Cell-3
Cell-4
GSECharge &Dischrg +
GSE Full Batt V Mon+
GSE Cell 1 V Mon+
GSE Cell 1Temp Mon
GSE Cell 2 Temp Mon
GSE Cell 3 Temp Mon
GSE Cell 4 Temp Mon
GSE Cell 2 V Mon+
GSE Cell 3 V Mon+
GSE Cell 4 V Mon+
GSE Cell 1 V Mon -
GSE Cell 2 V Mon -
GSE Cell 3 V Mon -
GSE Cell 4 V Mon -
GSECharge &Dischrg -
GSE Full Batt V Mon -
Battery Power +
Battery Power -
Full Batt V Mon -
Cell 4 Temp Mon
Cell 3 Temp Mon
Cell 2 Temp Mon
Cell 1Temp Mon
Full Batt V Mon+
Battery Schematic
Lithium-Polymer cellsPrismatic design4 A-hr capacity4 cells in series4 flight temp monitors4 GSE temp monitorsCell voltage to GSEBattery voltage to flight
Battery
Credit: Valence Electronicswww.valence.com
Back to Index
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DINO PDR April 21, 2023
Solar Array CIGS Cell Interconnects
5 Bonded CIGS PV on Stainless Steel Substrate
Etched Structural Bond Area
Electrically Conductive Adhesive
High Strength Space Qualified Dielectric Adhesive
Masked Electrical Bond Area
Back to Index
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DINO PDR April 21, 2023
Blanket Configuration
X1 Cells
Y (8 Cells)
String
6 Modules interconnected by MSI using reinforced Kapton
tape to form a string of 48 cells. 2 strings per solar array
wing.Array-Wing
2 Strings mechanically joined by MSI using reinforced Kapton to form a wing – 0.552 m2
Back to Index
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Operations
Appendix J
Mechanisms
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Drag-along vs. Leave Behind
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Tether
• Low-density polyethylene tape– Along length are three strands of
0.003 inch thick Spectra® 1000• Used for rip-stop protection, not load-
bearing
• 0.005 inch by 1 inch cross-section• Linear Density = 0.002173 kg/m• Young’s Modulus = 3.069 x 1010
N/m2
• Flown on Advanced Tether Experiment (ATEx) mission
Tether Material
Back to Index
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Braking System
• Uses friction to dissipate deployment energy
• Brakes tip-mass in last 3 ft. of deployment
• Materials– Wheel – Anodized Al– Brake Shoe – Delrin AF
• Ediss = 8.22 in-lb
– Ff =.2284 lb
Brake shoe
Tether
Wheel
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Aerofins MountingCups
Cones
Back to Index
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Cup Cones
Back to Index
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Release System - HOP
• Pin Puller• Less then 120g• 50 lbs of force• One HOP releases
4 Deployables• Total travel of HOP release
Pin: .3in• Activated with 28V at 18 watts for 2 minutes,
which heats up the wax inside the piston, expanding it and causing the pin puller to move
• HOP releases rings that attach to the release system via steel cable• Simultaneous release of Aerofins
Back to Index
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Aerofin Release
Ring Force (lb) Coefficent of Static Friction Coefficient of Kinetic Friction Factor of Safety Min Stowed Pull Force (lb) Min Deployed Pull Force (lb)30 0.2 0.15 2 12 9
Number of Springs Stowed Force per Spring (lb) Deployed Force per Spring (lb)2 6 4.5
Stroke Length (in) Min K Value Min Stowed Compression (in) Min Deployed Compression (in)0.25 6 1 0.75
Aerofins
Back to Index
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HOP Release
Back to Index
HOP
Ring Forces (lb)
Coefficent of Static Friction
Coefficient of Kinetic Friction
Factor of Safety
Min Pull Force (lb) (initial)
Min Pull Force (lb) (Final)
12 0.2 0.15 2 9.6 7.2
12
Total (lb) Stroke Length (in)
24 0.31
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Operations
Appendix K
Software
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FSW Concept of Operations
The Flight Software (FSW) will take incoming commands and perform the task requested (i.e. getting sensor readings, processing images, sending files or other calculations).
• A command will be posted on the Software Bus (i.e. a system of message queues) by a user or by another process.
• All processes listening for this particular command will pick up a copy of the command and perform the task associated with that command.
• If the process has to talk with a piece of hardware, it will send the appropriate hardware command to the appropriate driver.
• If the process expects a response before proceeding, it will wait, otherwise it will continue with its tasks.
• The FSW is low level code that performs the tasks it is given. When done, it will return a status to the calling function.
Back to Index
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SCL Concept of Operations
The SCL Real Time Engine (RTE) will perform the high level mission operations (i.e. decisions of which tasks to run, when to retry a task, when events are supposed to happen, etc.)
• SCL will ask the FSW to obtain sensor readings (Health and Status) periodically throughout the mission. This data is sent to SCL via a software bus message and stored in the database.
• As sensors change in the database, rules will watch the values and act appropriately.
• If a schedule of events is provided by Mission Operators, this schedule is carried out at the prescribed times.
• If an event is scheduled to occur and SCL decides that there are insufficient resources to perform this task (i.e. not enough power, not enough daylight, hardware not available, etc.), SCL will choose a new time to perform this task based upon built in heuristic methods.
• If a science event indicates that a follow-up event should happen, this unscheduled opportunity will be acted upon within SCL.
Back to Index
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VMOCC Concept of Operations
VMOCC is a product already developed for use on two other satellite systems. It is a group of software products that is used to get data to and from the spacecraft. The software was designed to satisfy future missions such as DINO.
• Commands are received from users worldwide via a webpage.• Commands are concatenated into a schedule.• The schedule is sent to the spacecraft via the most appropriate
ground station.• Telemetry from the spacecraft is received via a trusted or public
ground site.• The telemetry requested from the spacecraft is forwarded back to
the user via a webpage.
Back to Index
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Detailed Review (FSW)
• ProcessTask (FSW base class)– Receives cmd within the CMDBLK from route_cmd.– Initial processing drops into the Process_Cmd
function.– Responses are sent to and received from hardware
through IOBLK.– Processing a response from hardware is handled
though Process_Reply– Returning a status for each command is through
Process_Reply.Back to Index
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Detailed Review (FSW cont)
• Hardware I/O managers (i.e. serialmgr, I2Cmgr, usbmgr, pcmciamgr, 802mgr)– All hardware I/O managers inherit a base IOMGR
class• Data passed to hardware through odata• Data passed from hardware through idata
– Wireless 802.11 devices will have to inherit pcmciamgr
– Hardware I/O managers implement the driver specific for their hardware
– Protocol for talking to hardware could be uni-directional or bi-directional. Back to
Index
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Detailed Review (FSW cont)
• Communication protocol– TCP/IP based
• Satellite will have an IP address• Satellite will have password/firewall security
– Use standard telnet and ftp daemons• A broken FTP upload or download can be resumed on next
pass• Standard UDP broadcasts (H&S) can be received by any
computer with a radio
– Use built-in socket connections for SCL
Back to Index
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Detailed Review (FSW cont)
• Science analysis - Cloud height algorithm– Find common points on two images
• One is a height reference• One is a rotation reference
– Triangulate the pixel changes– Notify SCL if we image a significant cloud formation
• Science analysis - Topo map algorithm– Cloud heights taken on grid points are combined into
file– Interpolation between points may be possible Back to
Index
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Detailed Review (SCL)
• SCL (Overview)– A command comes in via a socket– The RTE decides what to do with it: immediate
command, run a script, run a FSW command– The SCL model contains:
• Scripts to execute mission objectives• Rules to fix problems
– The database contains records that hold sensor and derived information
– Telemetry is gathered from the database and sent to the ground
Back to Index
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Detailed Review (SCL cont)
• SCL model (DINO model needed)– Scripts to execute mission objectives
• The ICD from each subsystem will enable us to create scripts to run their components
• Software needs commands, parameter, and timing information
– Rules to fix problems• The System Team will tell us what can be done if a sensor goes out
of limits
– Database definition• Each sensors from every subsystem will have an entry
• Each value we want to calculate (derive) will have an entry
• Each piece of hardware will have a field to store operational statusBack to Index
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Detailed Review (VMOCC)
• VMOCC (overview)– Get commands from users worldwide– Deliver archived data (if it satisfies request)– Create a schedule for the satellite– Automatically initialize a radio (anywhere in the world)
and send the command– Accept telemetry from any radio in the world and
integrate into VMOCC model– Deliver requested data back to users
Back to Index
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Detailed Review (VMOCC cont)
• VMOCC (changes needed)– Use TCP/IP protocol similar to CX (but without PPP)– Ground database (MySQL) needed for telemetry
being sent– Ground SCL model (i.e. ground scripts and ground
rules) to alert mission operators of problems– STK model needed to get communication
opportunities, science data, and an attitude visualization
– Web interface (i.e. webpage) needed for commanding and receiving data
Back to Index
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Software Data Mang. and Testing
• Data Management and Testing– CVS for version control (i.e. code identification)
• archive code - protects against losing capabilities
• allows multiple users
– CVS will tag a baseline Release • the release will always be available to anyone authorized to use it
• programmers can directly go from one release to work on another
• testers can go back to previous release to verify capabilities still exist
– Testers job to account and audit what goes into a Release • Regression test all capabilities prior to each Release
• Tester's sign-off sheet to document functionality of each capability
• list all modules that need to be included
– All tests to be completed and satisfactory prior to each Release release • if not, decide on de-scope or delay options per Release
• all documentation to be on-line
– CVS to control access to files under CM
DINO PDR April 21, 2023
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Operations
Appendix L
Structures
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Main Structure
• Mass: 5.54kg• Height:12.5in • Diameter: 17.76in• Material: Al 6061• Similar design to Three
Corner Satellite
Back to Index
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Nadir and Zenith Plate
• Iso-Grid Design• Mounting plate for
Lightband systems
Back to Index
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Side Panels
Back to Index
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Camera Box Pieces
Outer Box-Holds camera at 30 deg angle -Allows access to USB port and power supply-Protects circuit board-Dimensions: 1.50 in. X 4.07 in.- 1/8 in. thick walls
Camera Bracket
-Secures camera in box
-Supports circuit board
Back to Index
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Camera Box Pieces Con.
Mounting Plate-Adapts to mounting holes in
structure
-Secures to Outer Box
-Allows for camera lens to look out of the box
Back to Index
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Complete Camera Box
Full assembly of camera box
Complete Assembly-Encloses camera
-Mounts to earth facing plate
-Holds camera at 30deg angle
-Manufacturing done as a short component
Back to Index
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Operations
Appendix M
Tip Mass
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TM-EPS Block Diagram
3.7V, 4A-hrLi-MP
Cell
5V DC/DCConverter
Inhib2
Inhib3
Inhib1 Sub-
Systems
Sep Sw.#1
Sep Sw.#2
PhotosensorOn/off sw.
System linesInhibit lines
GroundBack to Index
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Features - WISER 2400
• Operates In the ISM band (2.4GHz – 2.495GHz), no FCC license required
• No driver on the host device is required for radio operation
• Radio operation is independent of the operating system on the host equipment or device as long as a RS232 port is properly supported
• Industry standard IEEE 802.11b-compliant wireless interface; Interoperable Client radios from other vendors ( in our case a Wireless LAN card)
• Plug and play device; Once configured using a utility software, the configuration settings and other information is stored in non-volatile memory
Back to Index
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Specifications-WISER 2400
• Frequency: ISM band (2.4GHz – 2.495GHz) • Link Distance: ~1200 ft in open space • Voltage, current: 5v, max 480mA (in transmit mode)• Data rate: Capable of supporting up to 115K baud
(possible limitation on the digital camera side to transmit data)
• Weight : 3.7ounces:the radio with case,1.7 ounces is the weight of the case
• Antenna type: Integrated dipole antenna (omnidirectional) with ~2dBi gain
Back to Index
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Power Safety and OperationsSafety• Two-fault tolerant battery inhibit
system • No shunt diodes• Battery case must contain any
leaks and prevent shorts in the battery
• Fuse must be provided on ground leg of battery
• System will be un-powered until TM separation from DINO– Must be able to launch with a
charged battery• If Ball cannot help, will have to
use a multi-cell NiCd system– Need shunt diodes on ea. cell– All cell vents must be oriented
upward during launch
Operations• System will switch on and off via a
photo-sensor
– C&DH will turn the system back off if no images are to be taken on a given orbit
• Inhibit switches open until the tether is deployed
– Separation switches detect tether deploy
Back to Index
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Trade Study of Tip-Mass Camera
Fuji MX-1200 Samsung 800KKodak DC3200
Price $141.36 (Refurbished)
$143 (New) $209.99 (New)
Connection RS-232 RS-232 RS-232
Power 5V DC Port 5V DC Port 5V DC Port
Dimension 4.3 x 3 x 1.3 3.3 x 3.1 x 1.25 4.45 x 3.1 x 2.1
Weight 200 grams 190 grams 215 grams
Memory 4 MB SM Card 2 MB SM Card 2 MB(internal)
Shutter Speed
½ to 1/750 sec ½ to 1/1000 sec ¼ to 1/500 sec
Back to Index
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General Statistics Table
Component Mass (± x) Box Size Power Needed Current Draw Duty Cycles
Comm Unit TBD TBD 5 Volts497mA – TX
TBDmA – RX
TBD
100%
Camera 68 g 1.625” x 1.8” x 5” 5 Volts ~150mA 50%
Batteries 500 - 600g 1.25” x 4.25” x 4.25” 0 voltsTBD N/A
Power Conditioning Equipment
80g 2” x 2” x 0.5” 3.7 Volts 100mA 100%
Tether Unit1.01 kg 3.5” x 4.0” x 4.0”
0 volts N/A N/A
TOTAL 5.0 kg 3.9” x 9.0” x 9.0” 4 Amp hours N/A N/A
Back to Index
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Operations
Appendix N
Thermal
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THERMISTORS
• Where are we going to put them?– They will be put on
sensitive parts of the satellite
• How many do we need?– We’ll need about 27– Could change depending
on placement of items and if items are added
Power 9 (4 for GSE)
Structures 5
Science 3
ADCS 4
C&DH 2
Comm 2
Tip-mass 2Back to Index
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THERMISTORS
• What will we use?– Probably Getting it from BCcomponents– Between 10K ohm and 100K ohm
• High resistance to draw less current
– Temperature range of -40 to 125 oC• Fits within required temperature range
– Costs between $0.50 and $1.50 each
Back to Index
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PHYSICAL MODEL
• The first step in our model is establishing and defining nodes:
Back to Index
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MODELING • SUPVIEW
– Each of the nodes are defined by corners which are described with respect to a reference point currently selected as the absolute middle of the DINO satellite
– This is then processed by a program called SUPVIEW provided by Bob Pulley at Ball Aerospace.
– By the end of the modeling process this model will be run for both the inside and the outside of the satellite and each of the compartments it contains
– Our error is generally about 10^ -6
Nodes and Corners
SUPVIEW
View Factors
Nodal Dimensions
Fluxes
Back to Index
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MODELING
• FINDB6– This program simulates its orbit in efforts to
establish how much energy each node is exposed to due to:
SUN, EARTH IR, EARTH ALBEDO
– This is where the coldest and hottest cases are established with respect to Inclination range.
Orbit Starting Day (Spring Solstice)Orbit Ending Day (1 year later)
Universal time of launch (3600 Seconds)Altitude (350 km)Inclination (0-51)
Period (1.525 hours)
ASSUMPTIONS:
FINDB6
INCIDENT ANGLES FROM
SUN AND EARTHBack to Index
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MODELING
• ALBEDO– Knowing incident angles of sun
and earth on the satellite we can then estimate the amount of energy hitting each surface
– This is where we establish two separate models for the absolute hottest and coldest orbits our satellite can be a part of, which is based on beta angles
– Turns out the hottest case possible is actually a fully sunlit orbit, while the coldest is the case where the satellite spends most of its orbit behind the earth.
ASSUMPTIONS:Solar and Earthshine Constants
Beta anglesOrbital Positions of InterestSUPVIEW, FINDB6 outputs
ALBEDO
Energy Fluxes per node
per orbital position Back to Index
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MODELING
• REFLECT– Knowing how much
energy each node is exposed to, it is essential to determine how much of this energy is absorbed and how much is emitted out of the material.
Node by node energyEmmited and absorbed
FINDB6
Energy EmmittedEnergy Absorbed
Back to Index
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MODELING
• TAK III– This is the final program that
will predict temperatures for each node at any time during the orbit
– All information from previous programs are interfaced here into one model
– Node to Node conductance values are calculated based on heat capacity and dimensions of each material represented by nodes
– The internal model is combined with the external model however there are still two separate models for the hot and cold cases
Back to Index
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PRELIMINARY RESUTLS
• NODAL TEMP. RANGES FOR HOT CASE– With information presented
today an internal model will be created which will give more accurate values and we’ll see that these ranges will change significantly
NODES MIN MAX
Top Side -76 C 20 C
Bottom Side -76 C 20 C
Earth Face -80 C 20 C
Sun Face -80 C 20 C
Leading Edge -77 C 20 C
Airfoils (outside) -80 C 20 C
Airfoils (inside) -80 C 20 C
Trailing Edge -80 C 20 CBack to Index