snap spacecraft orbit design
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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 - PowerPoint PPT PresentationTRANSCRIPT
SNAP Spacecraft Orbit DesignSNAP Spacecraft Orbit Design
Stanford University
Matthew Peet
Presentation LayoutPresentation Layout
Mission requirementsThe use of swingby trajectoriesPrevious researchResearch goalsStatus of current workPlans for future work
SNAP Mission RequirementsSNAP 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 TypesCandidate Orbit Types
Low Earth Orbit Geostationary Orbit
High Earth Orbit Lagrange Orbit
Preferred Orbit DesignPreferred 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 RequirementsLaunch 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 AssistsGravity 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
Swingby trajectoriesSwingby 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 ResearchPrevious 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 DesignInterplanetary 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 ConceptPrometheus 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 AnalysisSatellite 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 ResearchGoals of Current Research
1. Reduce launch costs by minimizing the delta-v required to place the SNAP satellite in its optimal orbit
2. Facilitate mission planning by developing an analytic process that will produce an optimal lunar assist trajectory given launch date and desired orbit
3. Improve the analytical process to provide long-term orbit stability
Status of Current ResearchStatus 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 TrajectoryBaseline 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 TrajectoryDevelopment 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
Baseline TrajectoryBaseline Trajectory
Final Orbital Elements:– Rp = 11 Re– e = .696– i = 55.3 deg– RAAN= 354.3 deg– AOP = 22.3
Development of Analytic MethodDevelopment of Analytic Method
Consists of 3 stages
–Intercept stage
–Swingby stage
–Return stage
–Intercept stage
–Swingby stage
–Return stage
Intercept StageIntercept 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 StageDevelopment 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
Swingby StageSwingby Stage
Within sphere of influence, use simplified 2 body orbital motion
Relate exit conditions to arrival conditions
Development of Swingby StageDevelopment 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
Return StageReturn 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 StageDevelopment 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
Improved Baseline TrajectoryImproved Baseline Trajectory
Improved orbital characteristicsRp = 20 Ree = .399i = 73 degRAAN = 351 degAOP = 221.5 deg
High InclinationHigh Inclination
Inclination of 73 degrees– Reduced eclipse time to 5.6 hours– Only 82 minutes in the umbra
Orbital StabilityOrbital 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
Orbit StabilityOrbit Stability
Coverage TimeCoverage 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 CostsLaunch 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 WorkPlans for Future Work
Orbital Stability InvestigationImprove Matlab modelsDesign semi-analytic tools similar to
the Ocampo research