Space Based Solar Power Satellite Conceptual Design for Retrodirective Control

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Space Based Solar Power Satellite Conceptual
Design for Retrodirective Control

• Space Engineering Institute
• Spring 2009

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Overview

• Creating a satellite module that will be attached
  to a Japanese experimental satellite in a
  Low-earth orbit.
• Our team objective is to create a satellite
  module that can test the retrodirective beam
  control method of sending microwave power back to
  Earth.
• The module must provide its own power and have
  its own thermal management systems.

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Module System Sandwich Design
Space Environment Orbit
Photovoltaic Cells
Structures Materials
Thermal Management
Energy Storage Power Conversion
Antenna Array
Retrodirective Control logic
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Environment Analysis
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Semester Goals and Expectations

• Become familiar with Satellite Tool Kit software
• Model a cubic satellite and evaluate solar energy
  collected at each of its six faces to identify
  the optimum location for the solar panels
• Determine the maximum minimum and average solar
  energy collected on the optimum location during
  one orbit

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Satellite Orientation

• Orange-North
• Green- South
• White- Nadir
• Yellow-Zenith
• Purple-Leading
• Teal- Trailing

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Satellite Orbit Options

• Geostationary
• Altitude 35786 km
• Inclination 0º
• Low Earth Orbits
• Critically Inclined Sun Synchronous
• Perigee altitude 400 km
• Retrograde inclination 116.565º
• Circular
• Altitude 500 km
• Inclination 45º
• ETS-VII Japanese satellite with similar initial
  conditions
• Altitude 550 km
• Inclination 35º

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Energy and Power Received

• Power a cos()
• angle between the sun vector
• and the vector pointing normal to
• the face
• Units W/m2
• EnergyPowerTime
• Units J/m2
• Also dependent on the cosine of

http//solar.mridkash.com/wp-content/uploads/cosin
e-law.jpg
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Approach

• STK provides angular data for each face of the
  cubic satellite
• Angular data are converted to power W/m2
• Power data are converted to energy data J/m2

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Data from STK
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Data from STK
The Zenith Nadir Leading and Trailing faces
have approximately the same exposure to the sun.
The antenna will be located on the Nadir face so
it reasonably follows that the solar panels be
placed on the Zenith face directly opposite the
antenna. A circular orbit with an altitude of 500
km and inclination of 45 was chosen because
with the options available this orbit allows the
solar panels to receive the most sunlight.
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Thermal Management
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Thermal Management

• Objective
• To perform thermal analysis of the satellite and
  ensure a suitable operating environment for the
  payload.
• Tools
• Thermal Desktop software
• Research Topics
• Low earth orbit environment
• Temperature requirements for internal components
• Cooling/heating methods

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External Environment

• In LEO the satellite will be heated by
• Direct sunlight
• Earths albedo
• Earths IR emittance
• The total heat absorbed by the satellite will not
  remain
• constant. Fluctuations occur due to
• Entering/exiting Earths shadow
• Varying surface conditions on Earth

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Satellite Interior

• The interior environment of the satellite must be
  kept at
• a proper temperature range. Most electronic
  equipment
• onboard must operate in a surrounding temperature
• range of 0 to 50 degrees Celsius.
• Factors to consider for the internal energy
  balance
• Fluctuating external heat rates
• Heat released by electronic equipment
• -Low level baseline operation
• -High level during periodic transmission
• Thermophysical properties of structural material

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Cooling/Heating Methods

• External
• Radiators Do not require energy. Release heat
  without re-entry (thermal diode)
• Internal
• Thermoelectric Coolers/Heaters Require energy.
  Can absorb/emit heat by reversing polarity
• Mechanical cooling Expander compressor or heat
  exchanger. Takes up space and weight.
• Resistive Heating Requires energy but elements
  are compact in size.
• Heat Pipes Passive

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Thermal Desktop

• Objectives
• Develop a model for the satellite module.
• Use the orbital information from STK to determine
  thermal environment of the satellite.
• Progress
• In process of creating models.

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Thermal Desktop Continued
Example of absorbed flux from sun earths albedo
and IR emittance.
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Materials and Structures
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Structural Requirements

• The satellite must have ability to
• Withstand launch loads
• Provide desired rigidity
• Protect sensitive payload components from extreme
  temperatures.

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Material Selection
Currently evaluating two different materials

• Ti6Al4V Titanium alloy VS. Aluminum Alloy(
  7075-T651)
• Although Titanium is 60 heavier than Aluminum
  it is over twice as strong.
• Possibility of having titanium based honey comb
  exterior joined by a smaller portion of aluminum
  interior.

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Weight Comparison
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Honeycomb Layer

• Planned use of Honeycomb design
• Hexagonal Structure
• Uses the least amount of material to create a
  lattice of cells within a given volume
• Maintains strength

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Preliminary Sandwich Structure
Layered design that takes advantage of each
materials different thermal properties.
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Energy System
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Goal Requirements

• Collect Solar Energy and store it to power the RF
  amplifiers
• Collect power from 1m2 solar panel.
• Store energy in a medium that can withstand high
  drain current.
• Energy storage mediums must have a wide operating
  temperature range.

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DC to RF Converter Options

• Tube magnetron _at_ 5.8 GHz
• Can output 650W with 65 efficiency
• Heavier than solid state options
• (1.1kg vs .6g)
• Produces more heat than solid state converters
• Requires a high voltage power supply
• to excite the electrons.
• GaN HEMT solid state converters _at_ 5.8 GHz
• Fujitsu converter can output 320W theoretically
• Cree converter can output 35W commercially
• available now.
• Lightweight (.6 g) and extremely small size
• relative to the magnetron.

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Energy Storage Li-ion

• For storing energy from the photovoltaic cells
  Li-ion and Li-S batteries are being considered.
• Li-ion batteries have an energy density of 110
  Wh/kg.
• Saft MPS space series batteries that are already
  thermally insulated and autonomously heated.
• Have a wide operating temperature range ( -5o F
  to 140o F for charging and -40o F to 140o F in
  operation)
• Built in over current and charging circuits into
  the module.
• 17 Ah capacity per battery _at_ 28V.

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Energy Storage - Continued

• Lithium sulphur batteries are being considered
  for their higher energy density (350 Wh/kg vs the
  110Wh/kg for Li-Ion)
• Experimental expensive technology.
• No history of satellite use.
• Ultracapacitors
• High energy density capacitor used for powering
  the Solid state microwave converters when
  transmitting a signal.
• Ultracapacitors can handle 20A continuous
  current.
• Will be used in conjunction with the Li-ion
  batteries to power the GaN HEMT amplifiers at
  their maximum capacity

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Current Concept

• Maximum Power Point Tracker monitors the voltage
  and current of the Solar Panel and tracks the
  peak point on the power curve.
• Battery Management System tracks the charge rate
  voltage and current.

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Antenna and Retrodirective Control
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Retrodirective Beam Control

• The implementation of retrodirective beam control
  is critical to accurate beam pointing as well as
  the overall safety of the system.
• The key objective is to have the power beam of
  the solar power satellites transmitter pointed
  only in the direction of a received pilot beam
  which provides a phase reference
• Retrodirective beam control ensures that
  microwave power transmission is both safe and
  insusceptible to accidental misalignment.

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Proposed Retrodirectivity Method

• The 2.9 GHz incoming pilot signal is received at
  a Frequency of 1 and Phase f1
• To conjugate the received signal is next mixed
  with a reference source of Frequency 21 and
  Phase fref
• The conjugated signal is then mixed itself to
  produce a signal with Frequency 21 and Phase
  -2f1
• After conjugated and doubled the signal is
  transmitted from a different transmitting
  subarray
• The complete phased array transmits a 5.8GHz beam
  in the direction of the incoming pilot signal

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Proposed Phased Array Antenna Concept

• Linear Microstrip Patch Phased Array Antenna
• The Microstrip Patch Antenna will operate at a
  Frequency of 5.8 GHz and will have an Input
  Impedance of 50
• The Antennas design features a 4x4 Phased Array
  consisting of 15 Transmitting Elements and 1
  nested Receiving Element each spaced 0.5 apart
• The 4 subarrays are expected to be at different
  phases prior to power transmission

5.8 GHz 4x4 Linear Microstrip Patch Phased Array
with nested 2.9 GHz Receiving Element
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Antenna and Transmitter Interface

• Solid-State
• Facilitates electronic beam steering
• Power amplifier and phase shifter are placed
  behind each transmitting element
• Microwave filters are required to countervail
  amplifier-spawned noise

• Magnetron
• RF power is split to feed fed to each antenna
  subarray
• Negligible power loss may occur during energy
  feed
• Loss expected from phase shifter

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Advantages of Proposals

• Microstrip Patch Antenna
• Advantages
• Low cost to manufacture
• Light weight and low profile
• Supports both Linear and Circular Polarization

• Retrodirectivity Method
• Advantages
• Conjugates pilot signal directly at RF
• Reduction in the number of electronics per
  antenna subarray
• Less power consumption