coefromsv_六根数_轨道六根数_六根数轨道_卫星轨道_whole2sy

function [r,v] = sv_from_coe(coe) % %This function computes the state vector (r,v) from the %classical orbital elements(coe). % %mu-gravitational parameter(km^3;s^2) %coe-orbital elements[h e RA incl w TA] % where % h=angular momentum(km^2/s) % e=eccentricity % RA=right ascension of the ascending node(rad) %incl=inclination of the orbital(rad) %w=angument of perigee(rad) %TA=true anomaly(rad) %R3_w - Rotation matrix about the z-axis through the angle w %R1_i - Rotation matrix about the x-axis through the angle i %R3_W - Rotation matrix about the z-axis through the angle RA %Q_pX - Matrix of the transformation from perifocal to % geocentric equatorial frame %rp - position vector in the perifocal frame(km) %vp - velocity vector in the perifocal frame(km/s) %r - position vector in the geocentric equatorial frame % (km) %v - velocity vector in the geocentric equatorial frame % (km/s) % %User M-functions required:none % global mu h=ceo(1); e=ceo(2); RA=ceo(3); incl=ceo(4); w=ceo(5); TA=ceo(6); %...Equations 4.37and4.38(rp and vp are column vectors): rp=(h^2/mu)*(1/(1+e*cos(TA)))*(cos(TA)*[1;0;0]+sin(TA)*[0;1;0]); vp=(mu/h)*(-sin(TA)*[1;0;0]+(e+cos(TA))*[0;1;0]); %...Equation4.39: R3_W=[cos(RA) sin(RA) 0 -sin(RA) cos(RA) 0 0 0 1]; %...Equation4.40: R1_i=[1 0 0 0 cos(incl) sin(incl) 0 -sin(incl) cos(incl)]; %...Equation4.41: R3_w= [cos(w) sin(w) 0 -sin(w) cos(w) 0 0 0 1]; %...Equation4.44: Q_pX=R3_W'*R1_i'*R3_w'; %...Equations4.46(r and v are column vectors): r=Q_pX*rp; v=Q_pX*vp; %...Convert r and v into row vectors: r=r'; v=v'; %function [ output_args ] = Untitled3( input_args ) %UNTITLED3 此处显示有关此函数的摘要 % 此处显示详细说明