CL_ip_escapeDv

Description

• Computes the delta-V needed for the transfer from an elliptical orbit around a given planet to an hyperbolic escape orbit.

  The escape maneuver is tangential, i.e. the delta-V is parallel to the velocity vector on the elliptical orbit.

  The initial orbit is defined by its semi-major axis (sma) and eccentricity (ecc).

  The final orbit is defined by its velocity at infinity (vinf).

  The true anomaly of the maneuver on the initial orbit (tanoe) can optionally be specified. By default tanoe=0 (meaning 'at the periapsis').

  The planet is defined by its gravitational constant mu (default is %CL_mu)

Parameters

sma:

Semi-major axis of initial orbit [m]. (1xN)

ecc:

Eccentricity of initial orbit. (1xN)

vinf:

Target velocity on hyperbolic orbit at infinity [m/s]. (1xN)

tanoe:

(optional) True anomaly of maneuver [rad]. Default value is 0. (1xN)

mu:

(optional) Gravitational constant of the planet considered [m3/s2]. Default value is %CL_mu.

dv:

Norm of the delta-v. (1xN)

rph:

Periapsis radius of hyperbolic orbit [m]. (1xN)

tanoh:

True anomaly on the hyperbolic orbit at the time of the maneuver [rad]. (1xN)

Authors

• CNES - DCT/SB

Examples

// Escape from a LEO (circular) orbit : eqRad = 6378.e3; sma = [eqRad+350.e3 , eqRad+700.e3]; ecc = zeros(sma); vinf = [5 , 6] * 1.e3; [dv, rph, tanoh] = CL_ip_escapeDv(sma, ecc, vinf) // Escape from an orbit around Mars : eqRad = CL_dataGet("body.Mars.eqRad"); mu = CL_dataGet("body.Mars.mu"); sma = [eqRad+350.e3 , eqRad+700.e3]; ecc = [0.1 , 0.05]; vinf = [5 , 6] * 1.e3; tanoe = [0, 0.1]; // radians [dv,rph,tanoh] = CL_ip_escapeDv(sma,ecc,vinf,tanoe,mu)