Test the LPEccentric function. Provide the reference orbital elements, an 
   initial relative state and initial true anomaly, a desired/final relative
   state and final true anomaly, and the number of samples to discretize over.

   Two validation methods are used. In the first, we propagate the initial
   state over nS samples, applying the control acceleration found by
   LPEccentric. In the second, we integrate the two inertial states, applying 
   the same control acceleration, but in the ECI frame. We then transform
   back to the Hill's frame for comparison.

   Since version 7.
   [xH_int,xH_lin,dVTot,eT,err] = TestLPCircular( el0, xH0, xHF, dur, nS, plotFlag )

   el0             (6,1)  Initial orbital elements [a, i, W, w, e, M]
   xH0             (6,1)  Initial Hill's-frame state
   xHF             (6,1)  Final/desired Hill's-frame state
   dur              (1)   Maneuver duration [sec]
   nS               (1)   Number of samples to discretize over
   plotFlag         (1)   Create plots or not (0|1)

   xH_int          (6,1)  Hill's-frame state from integrating inertial states
   xH_lin          (6,1)  Hill's-frame state from propagating linear system
   dVTot            (1)   Total Delta-V for maneuver [m/s]
   eT               (1)   Time elapsed during call to LPEccentric [sec]
   err             (6,1)  State error at end of maneuver. Found by integrating
                          the inertial states from t=0 and applying control,
                          then comparing the final state with the target state.

	  Copyright 2004 Princeton Satellite Systems, Inc.
    All rights reserved.


FormationFlying: Dynamics/DiscreteHills
FormationFlying: Dynamics/FFIntegrate
FormationFlying: Dynamics/HillsEqns
FormationFlying: LP/LPCircular
FormationFlying: Transformation/Goals2Hills
FormationFlying: Transformation/Hills2ECI
Math: Linear/Mag
Orbit: OrbitCoord/El2Alfriend
OrbitMiniToolbox: Support/El2RV
SC: BasicOrbit/OrbRate
SC: BasicOrbit/Period