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SNAP Spacecraft Orbit Design. Stanford University Matthew Peet. Presentation Layout. Mission requirements The use of swingby trajectories Previous research Research goals Status of current work Plans for future work. SNAP Mission Requirements. Minimize Accelerations

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snap spacecraft orbit design

SNAP Spacecraft Orbit Design

Stanford University

Matthew Peet

presentation layout
Presentation Layout
  • Mission requirements
  • The use of swingby trajectories
  • Previous research
  • Research goals
  • Status of current work
  • Plans for future work
snap mission requirements
SNAP Mission Requirements
  • Minimize Accelerations
    • Improves target tracking
  • Minimize length of eclipse duration
    • Reduces onboard battery requirements
      • Weight(~1kg/kW-hr)
      • complexity
    • Heating and standby power reduced
  • Maximum contact with Berkeley
    • Allows increased data download
    • Improves control ability and reaction time
  • Avoid radiation belts
candidate orbit types
Candidate Orbit Types

Low Earth Orbit

Geostationary Orbit

High Earth Orbit

Lagrange Orbit

preferred orbit design
Preferred Orbit Design
  • High earth orbit
  • High inclination to avoid eclipse
    • >35 degrees required to avoid moon, but higher is better
  • Moderate eccentricity
    • Rp > 8 Re to avoid radiation
    • Ra < Rm to reduce antenna power
  • Apogee over northern hemisphere
launch requirements
Launch Requirements
  • For direct injection
  • three burns required for a total delta-v of 12 km/s
    • 1.75 km/s worth of fuel onboard for final burn
    • 2200 lbs of fuel for a 2000 lb spacecraft
  • Delta II Class Launch Vehicle Needed
    • Upper Stage Required
    • Cost: 80M-100M
gravity assists
Uses the gravitational attraction of a planetary body to alter the motion of a satellite.

Rotates relative spacecraft velocity in the planet-fixed reference frame about axes fixed to the planet.

Satellite energy is conserved within the planetary reference frame.

Planet-fixed frame is in motion with respect to the inertial space

A rotation in planetary system may not result in satellite energy conservation in inertial space

Gravity Assists
swingby trajectories
Swingby trajectories
  • Path of the spacecraft in planetary reference frame is rotated by angle delta
    • Sin(/2) = 1/e
    • e = 1 - Rp/a
    • a = 2*/v2
previous research
Previous Research
  • History of swingby trajectories in interplanetary mission design
    • Voyager, Pioneer, Magellan, Galileo, Cassini
  • Prometheus mission concept development
    • Long term observation strategy
  • Communications satellite rescue mission
    • Provided inclination change for stranded geostationary satellite
interplanetary mission design
Interplanetary Mission Design
  • First uses of Swingby concept
  • Restricted to in-plane maneuvers
    • No inclination changes
    • Allows for simplification
  • Voyager and Apollo through Cassini
prometheus mission concept
Prometheus Mission Concept
  • 1985 - current
  • First exploration of swingby trajectories for near-earth applications
    • Inclination changing
    • Perigee raising
  • Utilized a Monte-Carlo style technique
  • Never launched
satellite rescue mission analysis
Satellite Rescue Mission Analysis
  • 1998 - current
  • Development of technique for multiple passes
    • Insufficient fuel resources for direct encounter
  • Derivative based solution developed by Cesar Ocampo et al.
goals of current research
Goals of Current Research
  • Reduce launch costs by minimizing the delta-v required to place the SNAP satellite in its optimal orbit
  • Facilitate mission planning by developing an analytic process that will produce an optimal lunar assist trajectory given launch date and desired orbit
  • Improve the analytical process to provide long-term orbit stability
status of current research
Status of Current Research
  • Developed baseline trajectory based on adaptations of historical mission plans
  • Developed first order method for prediction and control of lunar encounter
  • Improved baseline trajectory based on analytical predictions
baseline trajectory
Baseline Trajectory
  • Launch: October 20, 2007
  • Based on Prometheus mission design
    • Earth observation satellite mission
  • Lunar intersection occurs at descending node
    • Eases adaptation of orbit
development of trajectory
Used STK with Astrogator to propagate orbit

Used 12th order earth model with perturbations out to 1/3 lunar distance

Runge-Kutta variable step propagator

Used 4th order selenocentric model with earth point mass and perturbations during lunar encounter

CisLunar variable step propagator

Used Initial trajectory identical to Prometheus mission

Calculated relative phase of moon in orbit at intersection during old mission

Calculated next occurrence for this phase starting in October, 2007

Determined launch date and time to intercept moon at this point in time

Development of Trajectory
baseline trajectory1
Baseline Trajectory
  • Final Orbital Elements:
    • Rp = 11 Re
    • e = .696
    • i = 55.3 deg
    • RAAN = 354.3 deg
    • AOP = 22.3
development of analytic method
Development of Analytic Method
  • Consists of 3 stages
  • Intercept stage
  • Intercept stage
  • Swingby stage
  • Swingby stage
  • Return stage
  • Return stage
intercept stage
Intercept Stage
  • Relate launch conditions to arrival conditions at moon
  • Find launch conditions for a given set of arrival conditions
  • Include effects of phasing loops and determine launch windows for desired conditions
development of intercept stage
Calculate launch conditions given launch date and azimuth

Calculate lunar position given intercept time

Apply phasing loops, if any

Propagate to lunar sphere of influence

Uses proportional error control to converge on solution

yields time of arrival and lunar position at arrival

Calculate relative position and velocity of the craft with respect to the moon at arrival

Given desired arrival conditions,relate back to specified launch conditions

Assumes constant arrival time at sphere of influence

work in progress

Development of Intercept Stage
swingby stage
Swingby Stage
  • Within sphere of influence, use simplified 2 body orbital motion
  • Relate exit conditions to arrival conditions
development of swingby stage
Translate relative position and velocity into Keplerian elements describing the lunar encounter

Propagate orbit through to edge of sphere of influence

Transform relative position and velocity to the inertial frame

Given beta-plane targeting parameters, calculate position and velocity at entrance to sphere of influence

Given exit position and velocity, determine beta-plane targeting parameters

The beta-plane parameters are used as outputs when the scenario is run through STK to ensure the values are roughly accurate

Development of Swingby Stage
return stage
Return Stage
  • Relate elements of final orbit to sphere of influence exit conditions
  • Assume an apogee lowering burn at perigee to provide orbital stability
development of return stage
Given position and velocity at edge of lunar sphere of influence, calculate new orbital element

Given new set of orbital elements, calculate apogee lowering burn for desired stability period

¼, ½, 2/3 lunar period, etc.

Find the final orbital elements following final burn

i and RAAN do not change

e can be related directly to elements at exit

efinal = 1-afinal(1-e)/a

Given desired Earth-Vehicle-Moon(EVM) angle and orbital parameters, determine initial AOP

not yet complete

Verify that desired orbital parameters meet Tisserand Criterion

Find exit position and velocity given desired orbital elements

entirely analytic solution

does not include mean or true anomaly

Development of Return Stage
improved baseline trajectory
Improved Baseline Trajectory
  • Improved orbital characteristics
  • Rp = 20 Re
  • e = .399
  • i = 73 deg
  • RAAN = 351 deg
  • AOP = 221.5 deg
high inclination
High Inclination
  • Inclination of 73 degrees
    • Reduced eclipse time to 5.6 hours
    • Only 82 minutes in the umbra
orbital stability
Orbital Stability
  • Three year nominal stability
  • Intrinsic stability of semi-major axis due to lunar influence
  • Slight reduction of inclination over lifetime of spacecraft
    • Increase in eclipse time is small
    • This stability issue will be explored in future work
coverage time
Coverage Time
  • Over the course of the three year lifetime
    • 60% is spent in Northern Hemisphere
    • 55.2% is spent in LOS contact with Bay Area
launch costs
Launch Costs
  • On-board fuel reserves require only 90 m/s
    • only 78 lb of fuel required
  • Launch Vehicle requirements reduced
    • C3 of –2 km^2/s^2
plans for future work
Plans for Future Work
  • Orbital Stability Investigation
  • Improve Matlab models
  • Design semi-analytic tools similar to the Ocampo research