MR SAT PROJECT University of MissouriRolla May 5, 2003

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MR SATPROJECTUniversity of Missouri-RollaMay
5, 2003
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MR SAT Overview

• Mission Objectives
• Design, produce, and launch a satellite for the
  advancement of satellites flying in formation
• Develop a low-cost communications link between
  spacecraft in the formation

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MR SAT Overview

• Why build a satellite?
• To mature key technologies for Distributed Space
  Systems missions
• To successfully integrate low-cost, innovative
  solutions while meeting mission objectives

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Mission Constraints

• Basic Requirements
• Operational life
• Minimum 2 weeks
• Goal 2 years
• Structure
• less than 45 x 45 x 45 (cm)
• Mass less than 20 kg
• Tether
• 10 meters
• Retractable

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Design Approach

• MR SAT Design Process
• Conceptual Design
• Research options, focus on most feasible approach
  to design
• Preliminary Design
• Integrate systems, begin prototype production and
  testing
• Final Design
• Design is locked in, fine tuning only

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Project Organization
Project Director
Systems Engineer
S.E. Assistant
Thermal
Power
Attitude
Orbit
Comm.
Propulsion
Structure
Launch Vehicle
Safety
Testing
Tether
Onboard Computer
Ground Station
SSE Lab
Web Site
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Miscellaneous

• Space Systems Engineering Lab
• Web Site (www.umr.edu/mrsat)
• Industry Mentors
• Fundraising

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Launch Vehicle, Safety, and Testing
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Launch Vehicle Goal

• To assure MR SAT can be carried into orbit by any
  available commercial launcher
• Physical constraints
• Testing and Safety requirements

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Possible Vehicles

• Space Shuttle
• Using the Shuttle Hitchhiker Experiment Launch
  System (SHELS)
• MR SAT constraints will be based on Shuttle
  requirements
• NASA Goddard may be able to help secure a launch

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Possible Vehicles
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Adapters
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Safety Goals

• Ensure safety compliance with launch vehicle
  systems
• Consult subsystems to avoid potential hazards
  during construction

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Testing

• Two modes of testing
• Functional
• Verifies that the components work to
  specifications
• Environmental
• Verifies that the components work under certain
  conditions
• Includes shock, shake, and thermal testing

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Testing Facilities

• Most tests can be performed at UMR campus
  facilities
• Off-campus facilities
• Boeing, St. Louis
• Kirtland AFB, New Mexico

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Current Status

• Several launch vehicle adapters are being
  designed
• Required documentation is still being collected
  and organized for future use
• Required testing results have been obtained
• Coordinating with several subsystems on testing
  requirements

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Structures
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Mission Constraints

• Total Mass Less than 20 kg
• Maximum Dimensions 45 x 45 x 45 cm
• Withstands Load Factor of 11g with a Factor of
  Safety of Two
• Natural Frequency of greater than 50 Hz
• Use low-cost solutions for structural components
  and construction

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External Shape
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Internal Mounting Surfaces

• Mounting Components
• Increase Strength and Rigidity
• Accessible for Ease of Construction

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Locking Mechanism

• Marman Clamp Design
• Controlled Detach
• Option to Reattach
• Power not Required to Maintain Lock

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Final Concept

• External Shape
• Internal Structure
• Solar Cells
• Tether Mechanism

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Joining Methods

• Friction Stir Welding
• High Strength Epoxy
• Fasteners

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Finite Element Analysis

• Static Analysis
• ANSYS
• 11g Loading
• Dynamic Analysis
• NASTRAN
• 50 Hz Natural Freq

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Current Status

• Performing Static and Dynamic FEA Analysis for
  Various Load Conditions
• Constructing Prototype Locking Mechanism
• Acquiring Materials for Structural Prototype

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Power
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Objectives

• Generate Power
• Store Power
• Manage Power

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Solar Cells

• Generate electrical power throughout mission
• Body mounted utilize all available surface area
• Triple junction gallium arsenide
• 24 efficiency
• 7 cm X 4 cm

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Solar Cells

• MR SAT
• 66 cells on each side
• One side to Sun
• Power produced - 60 W
• Two sides to Sun
• Power produced - 85 W

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Solar Cells

• MRS SAT
• 66 cells on top
• Top to Sun
• Power produced - 60 W
• Option 12 cells on each side
• One side to Sun
• Power produced - 10 W

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Power Required

• MR SAT
• Normal mode
• Computer, GPS, attitude control
• Total power required - 47 W
• Transmitting mode
• Total power required 72.5 W

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Power Required

• MRS SAT
• Normal mode
• Computer, GPS
• Total power required 22 W
• Transmitting mode
• Total power required 22.5 W

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Secondary Batteries

• Store electrical power throughout mission
• Possible options
• NiCd
• Li-Ion

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Power Distribution

• Distribute electrical power throughout mission
• Max power point tracker is used to regulate power
  generated by solar cells
• Charge batteries
• Power directly

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Power Distribution Schematic
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Power Distribution

• Voltage and amperage probes used on satellite
• Converters adjust voltage to needs of satellite
• Microcontrollers adjust power distribution to
  needs of satellite

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Current Status

• Solar panel prototype currently underway
• Energy from solar cells stored in batteries
• Size type yet to be determined
• Power distribution and regulation being designed
  to maximize efficiency
• Creating and analyzing duty cycle

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Battery Comparisons
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Power Budget
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Solar Cells Calculations
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Power Duty Cycle
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Propulsion
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Propulsion Subsystem

• Subsystem Purpose and Goals

Partial Orbital and Attitude Control of MR SAT
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Propulsion Subsystem

• Propulsion Mission Constraints
• MR SAT

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Satellite Propulsion Options Summary

• Mass and power restrictions ruled Out Ion and
  Hall Thrusters
• Safety requirements eliminated chemical rockets
  and hydrogen peroxide monopropellant
• MEMS eliminated due to low thrust production

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Current Options

• Primary Option
• Butane cold gas system
• Secondary Option
• Micro pulsed plasma thrusters

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Current Options

• Butane Thrusters
• Thrust levels 0.4N
• Simple construction
• Constructed from COTS parts
• Can be shelved for indefinite period
• Propellant can be loaded by students
• Orbit raising, attitude control, and chasing

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Current Options

• Micro Pulsed Plasma Thruster
• High specific impulse
• Orbit raising, attitude control, chasing
• Untested in space
• Thrust levels vary from uNs to mNs
• Power requirement issues

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Current Status

• Butane thruster designed
• Looking for parts
• Micro pulsed plasma thruster
• Prototype under construction

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Butane Thruster Drawing
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Orbit
MR SAT Tethered Satellite Project
University of Missouri-Rolla
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Mission Constraints
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Orbit

• Possible Orbits
• Circular
• Geosynchronous
• Sun Synchronous
• Low Earth Orbit

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Orbit

• Orbital Elements
• Based on Space Shuttle ISS service mission
• Inclination 51o
• Altitude 400 km
• Orbit Determination
• GPS

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Orbit

• Orbital Hazards
• Orbital Debris
• Van Allen Belts
• Atmospheric Drag

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Orbit

• Orbital Debris
• lt1 mm little to no hazard
• 1 mm to 1 cm possible damage
• gt1 cm damage to satellite will occur
• No way to actively track or avoid

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Orbit

• Van Allen Belts
• Begin at 640 km altitude
• Degrades satellite components
• Induces errors in digital circuits
• South Atlantic Anomaly
• Occurs because the Earths magnetic field is
  offset from its center.

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Orbit

• Atmospheric Drag
• Dependent on Sun and satellite cross-sectional
  area
• Eventual de-orbit of satellite

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Orbit

• Perturbations
• Earths Geopotential
• J2 Parameter
• Suns gravitational effect
• Moons gravitational effect

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Orbit

• Unknown Mission Constraints
• Launch Vehicle
• NASA missions
• Satellite avoidance
• Future Shuttle missions
• ISS orbit
• Hubble Space Telescope

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Attitude
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Attitude Subsystem

• Purpose Determine and Control Attitude of Both
  Satellites within Design Limits
• Subsystem Divisions
• Attitude Determination
• Attitude Control Devices
• Attitude Control Equations

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Mission Constraints
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Summary of Attitude Determination
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Global Positioning System (GPS)

• Uses relatively cheap receivers to process GPS
  signal and determine position and velocity
• Attitude determined using four antennas at a
  known distance apart
• High relative accuracy/price ratio
• Predicted accuracy dependent on available
  baseline possible range of 0.3 to 10 degrees

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Magnetometers

• Determine attitude measured relative to the
  Earths local magnetic field
• 0.3 to 1.2 kg mass, less than 1 W of power
• Predicted accuracy of 0.3 to 5.0 degrees
• Simple, reliable, lightweight, low-cost
• Uncertainties and variability of magnetic field
  dominate accuracy
• Daily cycling or otherwise predictable
  variations corrected for using sensor filtering

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Summary of Attitude Control
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Magnetic Torquers

• Currently in the design process
• Exploring the possibility of constructing in UMR
  Space Systems Engineering Lab
• Widely Used

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Cold Gas Thrusters

• Low Power Consumption, 1 W
• Exploring the possibility of constructing in UMR
  Space Systems Engineering Lab
• Widely Used

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Gravity Gradient Stability and Control

• Physics causes TSS to align so that it is
  oriented towardsEarths center with some
  libration in and out of the orbit plane
• Passive attitude control method
• Inherent aspect of tether dynamics - so we can
  either work with it or against it
• No power requirements, no lifetime limits, /-
  10 degrees accuracy

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Current Status

• GPS and Magnetometers provide determination
• Gravity Gradient, Magnetic Torquers, and Cold
  Gas Thrusters provide control
• Researching Vendors
• Exploring in-house construction possibilities
• Studying closed-loop equations

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Current Status, contd

• Testing a Magnetic Torquer Prototype
• Estimated to be able to turn the satellite
  through 180 degrees in 2.5 minutes
• 11 Watt power consumption

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Tether Dynamics
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Tether Dynamics

• Derive Equations of Motion (EOMs)
• Model and Simulate Dynamics
• Deliverables to Other Subsystems
• Current Status

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Derivation of EOMs for TSS

• EOMs for Tethered Satellite System (TSS)
  developed
• Complexities such as perturbations will be
  accounted for in subsequent derivations if
  necessary

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Model and Simulate Dynamics

• Use EOMs to model TSS dynamics
• Simulate dynamics with program using Runge-Kutta
  7-8th order numerical integration
• Verified core program using simple EOMs for
  pendulum motion test case and two-body problem

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Accurately Model Dynamics

• Comparison of our methods to other groups
  currently researching tether dynamics
• Intelligently balance assumptions, accuracy, and
  EOM complexity
• At a minimum, simulation accuracies must satisfy
  mission constraints for Orbit and Attitude
  subsystems

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Deliverables to Other Subsystems

• Max and min tether tension to Tether Structure
  subsystem
• Analysis of satellite operational modes including
  orbital injection, tether deployment, tether
  operation, severing the tether, and tether
  failure
• Verify simulation program for use by Orbit and
  Attitude subsystems

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Current Status

• Simulation program developed and verified using
  pendulum and simple two-body test cases
• TSS EOMs developed and arranged for use in
  simulation program
• Debugging program code
• Continuing to review available research in tether
  dynamics

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Tether Structure
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Tether StructureOverview

• Design aspects of the Tether Reeling Mechanism
• Proper tether material and diameter
• Control of tension in tether

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Tether Structure Mission Constraints

• Mass 2 kg
• Power 2 watts
• Design To deploy and retract 10 m tether
• Sensors to record tether performance

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Tether Structure
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Tether Structure
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Tether Structure

• Tether Deployment Retracting Sequence
• Sub-satellite given initial ejection velocity by
  springs
• Tether reels out due to ejection velocity
• Motorized Deployment Motor acts as a brake to
  control the rotational rate of the spool
• Tether can be reeled back in

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Tether Structure

• Tether Reeling Mechanism
• Reel
• Motor
• Guided or Pinch rollers
• Control unit
• Tether cutter
• 6. Auxillary Gears, Sensors

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Tether Structure

• Basic Design Parameters of TRM

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Tether Structure
Deployment Path

• Ejection Velocity
• assumed as1 m/min
• Minimal tether tension
• during deployment

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Tether Structure

• Current Status
• EOMs for the dynamic analysis of the TRM under
  progress
• Materials for the TRM under consideration
• Prototype of TRM to be built and tested
• Testing Kevlar and Dyneema tethers

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Onboard Computer
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Onboard Computer

• Overview
• Purpose
• Mission Constraints
• Proposed Design
• Current progress

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Onboard Computer

• Purpose
• Subsystem coordination
• Data storage and processing
• Communications control

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Onboard Computer
Mission Constraints

• MR SAT
• MRS SAT

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Onboard Computer

• Baseline Design

• Hybrid Design
• 386 CPU
• 8051 Micros

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Onboard Computer
Baseline Design

• Onboard Software
• 386 Single Board Computer
• Linux OS
• Custom Programs (C/C)
• 8051 Microcontrollers
• No OS
• Custom Programs (C/Assembly)

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Onboard Computer
Current Progress

• Hardware
• Single board computer
• 8051 interface
• Sensor integration
• Software
• Subsystem modes of operation

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Onboard Computer

• Current Progress

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Communications
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Mission Objectives

• Establish bi-directional Space-to-Ground
  Communications
• Establish bi-directional Inter-satellite
  Communications

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Mission Constraints
Space-to-Ground Communications

• Bi-directional communications
• Fast enough to transmit the desired data in a
  short time window
• Low power requirements
• ITU compliant
• NASA compliant

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Mission Constraints
Inter-Satellite Communications

• Affordable, and based on off-the-shelf wireless
  technologies
• Bi-directional communications
• Low power requirements
• ITU compliant
• NASA compliant

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Space-to-Ground Communications
Three technologies have been considered

• VHF-Band 145 to 900 MHz
• S-Band 2 to 4 GHz
• C-Band 4 to 8 GHz

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Inter-Satellite Communications
Three communication protocols have been
considered

• 802.11a
• 802.11b
• 802.11g

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Shielding
Two options are being considered to protect non
space rated equipment from radiation

• Aluminum Enclosures 5 mm thick
• Specific anti-radiation material developed by the
  RST company and named DEMRONTM

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Modes of Operation

• Off Mode
• Safe Mode
• Normal Mode
• Transmitting Mode
• Space-to-Ground Transmitting Mode
• Inter-Satellite Transmitting Mode
• Failure Mode

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Current Status

• Components have been selected and some purchased
• Next step start testing and integrating the
  selected components

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Thermal Analysis
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Introduction Objectives

• Satellite temperature normally governed by
  radiation
• Convection is negligible for the satellite
• Conduction depends on the heat transfer
  
  
  coefficient of the material
• Objectives
• Analyze satellite thermally
• Determine ways to make it safer
• Use of sensors
• Tether consideration
• Testing

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Analysis

• Software used is IDEAS/TMG.
• Provide passive thermal control for
• both satellites and their components
• Catalog of thermal properties of
• materials used

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Analysis Plan

• Lumped capacitance method
• Worst-case scenario
• Beta angle
• Cold case 00
• Hot case 900
• Steady state analysis
• Transient thermal analysis

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Lumped Capacitance

• Based on conservation of energy laws
• Power in Power out

Direct Solar Radiation
Radiation into Deep Space
Sun
MR SAT
Other Power losses - Radio Transmission Propulsion
Albedo
Earth Radiation
Earth
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Primary Analysis

• Minimum and maximum temperatures are
• -10 and 95 ºC respectively.
• Since the satellite was taken as a lump this is
  an average temperature and therefore there will
  be panels which will have lower or higher
  temperatures.
• This is why a detailed analysis is being
  conducted so that we can have a better
  understanding of what the temperatures in each
  panel will be.

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Optimizing Thermal Protection

• Either design to radiate or absorb
• Conductive links
• Conductive pastes
• Black paints on the inside surface
• Solar reflectors ( On the Bottom face of MR SAT
  )
• ( mirrors , white paints , Al backed teflon )
• Multiple layer insulation ( As shown )
• ( Probably Kapton)
• Cooling system

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Tether Consideration
Analyze temperature range for which tether
retains its properties
U.H.M.W Gel Spun ultra high molecular weight
polyethylene (Spectra, Dyneema) Percentages
mentioned in Ultraviolet exposure are result of
loss in strength due to one-year exposure to
sunlight

• Loss in strength due to exposure to sunlight
• Thermal expansion

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Testing

• T-Vac test
• Thermal cycle test
• Thermal Balance Test
• Sensor calibration
• Effect of orbital decay
• Critical temperatures for tether and mechanism
• Effect of variation in Beta angle
• Analysis of temperature vs. spin rate

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Current Status

• Steady state analysis and worst-case scenario
  using IDEAS/TMG analysis are under progress.
• Hardware aspects ( Multi-layer insulation, black
  paints, conductive pastes, sensors )
• Testing Options

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Conclusion
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Conclusion

• Current Status
• Conceptual design finalized
• Preliminary design in progress
• Mock-up in use
• Prototype structure planned

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Conclusion

• Current Challenges
• Funding
• Hardware
• Lab Development
• Graduate Students
• Mentor Recruiting
• Diversifying Team Disciplines (especially
  Electrical and Computer Engineers)
• Team Member Turnover

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Conclusion

• Upcoming Action Items
• Prototype construction to begin shortly
• Preliminary Design Document
• and Review
• Begin testing and integration of hardware and
  software

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Thank You!
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Send comments/questions to pernicka_at_umr.edu