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Orbital Debris Project

AERO 351 [Orbital Mechanics] Final project

Research 4 orbital debris objects and create a mission between the objects.

* This page is a summary of work done in a group. A lot of credit is given to my group members - my involvement is noted below.


 

Involvement

  • Lambert's Solution 
  • Transfer Optimization
  • Matlab Plots
  • Timelines
 

 

After 4 debris objects were found (1 GEO object, 1 MEO object, 2 LEO objects), two missions were created. The first scenario was required to take longer than 5 years with a total delta v  spent < 18 km/s. The second scenario was to finish within 5 years with a total delta v spent < 25 km/s. My group members worked on the second scenario while I worked on the first scenario - the second scenario is not included on this page for privacy. For both scenarios, the mission begins at one of the objects and finishes once the mission spacecraft rendezvouses with each object. For certain burns, impulsive maneuvers were assumed. The delta v required to leave earth's surface is not considered in the total delta v spent.

 

Debris Objects At Mission Departure

OBJECT ONE

  • LEO1 [Parking]
  • AAM/PSLV

COEs

  • Rp =  6,984.0 km
  • Ra =  7,026.8 km
  • Inclination =  5.95°
  • RAAN =  241.08°
  • Ecc =  0.0031
  • ⍵ =  8.5e-7°
  • a =  7,005.4 km

OBJECT TWO

  • LEO2
  • PEGASUS R/B 2012

COEs

  • Rp =  6,906.3 km
  • Ra =  6,929.3 km
  • Inclination =  1.96°
  • RAAN =  358.74°
  • Ecc  =  0.0017
  • ⍵ =  305.32°
  • a =  6,917.8 km

OBJECT THREE

  • MEO
  • ARIANE 5 DEB (SYLDA)

COEs

  • Rp =  6,596.2 km
  • Ra =  12,204.6 km
  • Inclination =  3.23°
  • RAAN =  132.36°
  • Ecc  =  0.2983
  • ⍵ =  267.01°
  • a =  9,400.4 km

 

OBJECT FOUR

  • GEO
  • JCSAT 6

COEs

  • Rp =  42,138.7 km
  • Ra =  42,191.1 km
  • Inclination =  2.57°
  • RAAN =  79.28°
  • Ecc  =  0.00062160
  • ⍵ =  289.97°
  • a =  42,164.8 km

 

 

Departure Date: 

November 2, 2017, 19h, 2m, 12s

 

 

Parking Orbit [LEO Object 1]

Indian Space Research Organization (ISRO), Polar Satellite Launch Vehicle

Launched February 21st 2015:

TLE

1 40937U 15052H   17313.38999385  .00000945  00000-0  33534-4 0  9995

2 40937   5.9543 241.0847 0030546 283.8112  75.8617 14.80647783114761

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Scenario 1 Lambert Optimization Scheme

An algorithm was written to find an optimal Lambert solution between two orbits.

 

  • Each color represents a different transfer leave time

  • Each line represents a different transfer time

  • Scheme determines transfer with least ΔV spent

  • If any orbit comes too close to Earth, solution is invalid

  • A wait period was implemented to find an optimal time to leave when synodic periods of the orbits lined up.

  • Used for all 3 transfers in scenario 1

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Transfers

3.jpg
rendermission_Legend.png

Feasibility of Mission

 

  • Total ΔV = 5.81 km/s < mission requirement.

  • Time spent on mission meets requirement.

  • Though a mission between LEO, MEO, and GEO is not necessarily practical in real applications, this mission is possible. The largest ΔV associated with completing such a mission is ignored by starting in a parking orbit. Getting into LEO from Earth’s surface costs ~9-10 km/s. Assuming the spacecraft isfully fueled in the parking orbit (LEO1), completing the mission is very feasible. Impulsive maneuvers are assumed which is not realistic, however, but a fair estimate for such a mission.

  • The 5 year time requirement is relatively long for a complete Lambert solution. An optimization function was written to run through thousands of scenarios throughout 5 years to find an ideal mission. Though compute time was a limiting factor when propagating for 5 years, an efficient solution was found.

Acknowledgements

My group

Dr. A