Gabe Karpati

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SuperNova / Acceleration Probe (SNAP)
System Overview
Gabe Karpati June 28, 2001
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Outline

• Driving Requirements and Assumptions
• Options
• Selected Configuration and Rationale
• Technologies Required
• Mass and Power Summary
• Requirements Verification
• Additional Trades
• Risk Assessment
• Issues and Concerns

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Overview

• Mission objective
• Determine the magnitude-redshift relationship for
  supernovae of type 1A over redshift range
  0.3ltZlt1.8
• Determine the distribution of gravitational
  potentials along cosmologically significant lines
  of sight
• Determine the magnitude-redshift relationship for
  other supernova types
• Additional constraints, challenges, and
  measurements
• Orbit w/ adequate thermal environment, good
  observing efficiency
• Pointing stability and tracking / Observatory
  stiffness
• Primary purpose of this study
• Establish/validate baseline mission configuration
• Length of study
• 4 days

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Driving Requirements Assumptions

• Orbit
• Best candidates to date are 19 Re x 57 Re w/
  lunar assist or 38 Re circular achieved w/ lunar
  assist
• Launch year 2008
• Lifetime 2 year required, 5 years goal
• Quality level Selective redundancy
• End-of-life disposal Not required for orbits gt
  GEO
• Instrument support
• Mass 700 kg
• Power 135 W max, 85 W avg, 22W stdby
• Average Instrument Data Rate 40 Mbps continuous

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Orbit Options

• LEO orbits
• PROs Simplify launch vehicle and propulsion
  requirements, easy RF comm
• CONs Enormous loss of observing efficiency, huge
  thermal problems for Payload, Bus, and Radiators
• Disposal required
• HEO orbits
• PROs Easy observing, good thermal environment
• CON Radiation environment problems
• SHEO orbits w/ lunar assist
• PROs Easy observing, good thermal environment
• CONs Requires exotic launch vehicle, difficult
  downlink and ground station situation that limits
  inclinations, orbits. Difficult orbit
  calculations.
• Sun-synchronous drift-away orbits
• PROs Perfect observing w/o eclipses, best
  thermal environment
• CONs RF communication difficult, especially at
  EOL. Possible mass constraint.
• Final orbit selection after detailed analyses.
  For some orbits, minor errors can easily
  propagate into fatal dispersion

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Launch Vehicle Options

• Delta 2920H-10L
• Liftoff capability marginal but fairing is way
  too small
• GSFC actual cost estimate is 67M to 72M
• Maiden flight w/ SIRTF in 2002
• Delta 3940-11
• Liftoff capability adequate
• GSFC actual cost estimate is 80M to 84M
• Liftoff capability adequate, fairing volume tight
  but adequate
• Atlas IIIB
• Liftoff capability comfortable, fairing volume
  tight but adequate
• Zenit 3SL w/ Block DM-SL restartable upper stage
  (Sea Launch)
• Liftoff capability comfortable, fairing volume
  comfortable
• For details see SNAP_Launch_Vehicles.xls

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RSDO Bus Options

• Ball Aerospace BCP 2000, Bus dry mass 608 kg
• Payload Power (OAV) (EOL) / Mass Limit 730 W /
  380 kg
• Spectrum Astro - SA 200HP, Bus dry mass 354 kg
• Payload Power (OAV) (EOL) / Mass Limit 650 W /
  666 kg
• Orbital StarBus, Bus dry mass 566 kg
• Payload Power (OAV) (EOL) / Mass Limit 550 W /
  200 kg
• Lockheed Martin - LM 900, Bus dry mass 492 kg
• Payload Power (OAV) (EOL) / Mass Limit 344 W /
  470 kg
• Orbital - Midstar, Bus dry mass 580 kg
• Payload Power (OAV) (EOL) / Mass Limit 327 W /
  780 kg
• For details see SNAP_ Candidate_RSDO_Busses.xls
• All above busses are designed for multi-year
  missions w/ redundant components.
• Mission Unique spacecraft structure and several
  significant subsystem upgrades are required.

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Other Options Considered

• Earth Shield for Detector Radiator
• Use in LEO to eliminate thermal coupling between
  Earth and Radiator
• May have to be actuated similar to a solar array
  drive
• Natural frequency must be gt 1Hz (preferably gtgt
  1Hz)
• Rigid cylinder shape could be considered
• PROs
• Would allow passive cooling of Radiator to 130K
  even at LEO
• CONs
• Extra mechanism, increased risk
• Complicates IT
• DISPOSITION
• Option dismissed, as LEO option was dropped due
  to several other problems

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Other Options Considered

• Slew rate vs. jitter
• Select small reaction wheels for low jitter or
  Heavy Duty (higher jitter) reaction wheels for
  fast slew
• DISPOSITION
• Analysis shows that the actual driver is solar
  wind, requiring bigger wheels
• 30 Nms/.05Nm wheels selected
• Use isolation mount

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Baseline Configuration Rationale

• 3-axis stabilized, 4 Reaction wheels, IRUs, no
  torquer bar
• Sun side w/ rigid body mounted solar arrays
  anti-sun side w/ radiators
• Standard Hydrazine propulsion system, 1
  lbs.thrusters, 150 kg total propellant required
  _at_ 100 m/s for apogee lowering, corrections and
  ACS.
• 74 Gbits SSR, storage only for spectroscopy data.
  (Avg. data rate 52 Mbps lossless compression
  plus overhead).
• Continuous Ka band downlink _at_ 55 Mbps to 3
  Northern Latitude ground stations (Berkeley,
  France, Japan).
• 3 gimbaled .7m Ka band HGAs S-band omnis S-Band
  TC thru omni antennas

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Preliminary Bus Subsystems Mass kg
Bus Structure 143.0 Payload
Mount 9.0 Antenna support 30.0 ACS 50.0
CDH 12.0 Power Electronics 15.0 Battery
81.2 Solar Arrays 10.5 Thermal
Hardware 67.0 RF Communications 53.0 Bus
Harness 8.0 Separation System, spacecraft
side 8.0 Bus Subsystems Total 486.7
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Preliminary SNAP Obesrvatory Mass kg
Payload Total 700 Bus Subsystems
Total 490 Propulsion Total (estimate) 170 SNA
P Observatory Total 1360
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Preliminary Cost Summary M

• Cost information
• Details in spreadsheet
• Subsystem cost estimates included
• Other costs are rough estimates

SNAP MISSION COST w/ Contingency SCIENCE
SPACECRAFT BUS MISSION
INTEGRATION OPERATIONS LAUNCH
VEHICLE TOTAL (excl. data analysis) 307.8M

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Requirements Verification

• Standard functional and environmental
  verification per GEVS
• Tight contamination control required
• Main challenges are in Instrument verification
• Ideally, observatory level thermal vacuum /
  thermal balance test is combined with and
  end-to-end image quality verification
• May use double-pass test

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Additional Trades to Consider

• Revisit possibility of using LEO
• Lower mission cost
• More realistic launch / launch vehicle
  configuration
• May eliminate propulsion
• Simplifies RF Comm
• Must assess disposal issues

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Risk Assessment, Technologies

• Risk on spacecraft bus is generally low, with
  well-understood technologies and readily
  available components.
• Designed subsystem for a 3 year mission, solar
  array sized to 5 years
• Higher risk on instrument, especially on the
  enormous CCD cluster
• No significant technology development required
  for bus

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Overview, Supporting Data

• Mass, Power, and Cost information
• SNAP_MassCost_Summary.xls
• Useful Web sites
• Access to Space at http//accesstospace.gsfc.nasa.
  gov/ provides launch vehicle performance
  information and other useful design data.
• Rapid Spacecraft Development Office at
  http//rsdo.gsfc.nasa.gov/ provides spacecraft
  bus studies and procurement services.