DINO PDR 7 October 2015 Deployment and Intelligent Nanosat Operations University of Colorado at...

DINO PDR April 21, 2023

Deployment and Intelligent Nanosat

Operations

University of Colorado at Boulder

University Nanosat III

August 14th-15th, 2003Logan, Utah

Colorado Space Grant Consortium 2


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



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.



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



Overall Mission TimelineLaunch

ICU DeploymentDINO Deployment

Mission Activation

Tether Deployment

FITS Deployment

Aerofin Deployment

Normal Science Operations

Detumble



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



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



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



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



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



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.



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



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



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



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



2. Frangibolt

Release Execution• Memory Composite • 500 lbs holding force• 25W @ 28V • 21 seconds• Increased temperature

from power activates release.



2. FITS Deployment



3. EMC Hinges and Release Mechanism


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



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



3. Aerofins Release Mechanism

StowedReleased



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



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.



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.



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



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.



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.



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



6. Electrical Power System

EPS Board



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%



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



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



ADCS Requirements


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



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



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



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



Software


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



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



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



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



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



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



Structures Requirements


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



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



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



Mass Budget

Tip-Mass Mass Budget

Comm8%

Science9%

Systems2%

Thermal2%

Power15%Str/Mech

39%

Margin25%


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



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



Power Budget

Tip-Mass

Power17%

Comm45%

Imaging11%

C&DH2%

Margin25%


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



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



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



Operations

Appendix



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



Operations

Appendix A

Requirements



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



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



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



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



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



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



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



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



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



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



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



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



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



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



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



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




– 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




• 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



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



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



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



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



Operations

Appendix B

Systems Charts



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




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)


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)


Back to Index




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)


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)


Back to Index



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



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



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



Power Budget

Normal Nighttime

Comm Tx

10%

ADCS11%

Comm Rx6%

Thermal1%

Power22%

C&DH25%

Margin25%


Subsystems

ADCS 11 1.0

C&DH 25 2.3

Comm 16 1.5


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



Power Budget

Safe Mode

C&DH31%

Power34%

Comm Rx9%

Thermal1%

Margin25%


Subsystems

ADCS 0 0.0

C&DH 31 3.3

Comm 9 1.0


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% 8 W-hr

Margin 25% 2 W-hr

Total 100% 10 W-hrBack to Index



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



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



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



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



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



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



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



Operations

Appendix CManagement and

Outreach



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



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



Schedule



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



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



Operations

Appendix D

Subsystem Block Diagrams



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



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



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



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



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



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



MOPS Flow Diagram

Sequence Execution

Sequence Adjustment

Command and Control

Data ProcessingFault Handling and Response

S/C Subsystems

commands

Raw DataBack to Index



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



RTE

S/C Model

Flow Charts - SCL

CMDIN TMOUT

database

Scripts

Incoming

data

Outgoing

data

Rules

FSW

Back to Index



Flow Charts – VMOCC Uplink

MySQL

STK

Disk/ Files

SCLRTE

TMOUT

Webbrowser

ScheduledCmds

Events

Schedules

Schedules

Cmds Pkts

= H/W Interface

Immediate CmdsBack to Index



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



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



Operations

Appendix E

Test Plans



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



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



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



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



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



FITS Release System – Test Plan

Gravity Back to Index



Aerofins – Test Plan

Aerofin Structure

Sled Air table

CTD Hinge

StructureAir table

Sled

Back to Index



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



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



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



Operations

Appendix F

Attitude and Determination Control System



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



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



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



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



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



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



Operations

Appendix G

Command and Data Handling



Power Consumption

PartPower Consumption

(Watts)

Flight Computer 3

Wireless Interface 3

Miscellaneous Interfaces 1(max)

Total 7(max)

Back to Index



RPX Lite

• SDRAM: 64M• FLASH: 16M• NVRAM: 128K• Real-Time Clock• Interfaces:

– USB

– RS-232 Serial

– PCMCIA

– Ethernet: 10BaseT

– SPI, I2C …

Back to Index



USB Interface

+5V

GND

D+

D-

Camera0

+5V

GND

D+

D-

Camera1

MUX

D+

D-

SelectVcc

Back to Index



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



Operations

Appendix H

Communications System



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



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



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



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



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



Operations

Appendix I

Electrical Power System



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



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



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



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



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



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



Operations

Appendix J

Mechanisms



Drag-along vs. Leave Behind

Back to Index



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



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

Back to Index



Aerofins MountingCups

Cones

Back to Index



Cup Cones

Back to Index



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



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



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



Operations

Appendix K

Software



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



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



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



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



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




• 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




• 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



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



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



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



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



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



Operations

Appendix L

Structures



Main Structure

• Mass: 5.54kg• Height:12.5in • Diameter: 17.76in• Material: Al 6061• Similar design to Three

Corner Satellite

Back to Index



Nadir and Zenith Plate

• Iso-Grid Design• Mounting plate for

Lightband systems

Back to Index



Side Panels

Back to Index



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



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



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



Operations

Appendix M

Tip Mass



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



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



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

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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

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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

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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

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Operations

Appendix N

Thermal



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



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

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PHYSICAL MODEL

• The first step in our model is establishing and defining nodes:

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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

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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



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



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

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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

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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

DINO PDR 7 October 2015 Deployment and Intelligent Nanosat Operations University of Colorado at...

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