A method of deorbiting a spacecraft, comprising: selecting, by a computing device of the spacecraft, a target landing site for deorbiting the spacecraft; during a predetermined time window associated with deorbiting the spacecraft: (i) determining, by the computing device, a range target and a velocity target for reaching a predicted atmospheric entry location, wherein the predicted atmospheric entry location corresponds to a point at which the spacecraft interacts with an atmosphere in order to reach the target landing site, (ii) determining, by the computing device, a back-propagated orbit state estimate of the spacecraft, wherein the back-propagated orbit state estimate of the spacecraft corresponds to an estimated location and path-angle of the spacecraft from the predicted atmospheric entry location based on the range target and the velocity target, (iii) comparing, by the computing device, the back-propagated orbit state estimate to a known orbit state of the spacecraft to determine that the back-propagated orbit state estimate has converged with the known orbit state, and (iv) calculating, by the computing device, based on determining that the back-propagated orbit state estimate has converged with the known orbit state, (a) an estimated time of ignition for a propulsion system of the spacecraft and (b) an estimated burn velocity vector of the propulsion system using the range target and the velocity target; and at or prior to an end time of the predetermined time window, performing a burn pulse by the propulsion system in accordance with the estimated time of ignition and the estimated burn velocity vector. selecting, by a computing device of the spacecraft, a target landing site for deorbiting the spacecraft; during a predetermined time window associated with deorbiting the spacecraft: (i) determining, by the computing device, a range target and a velocity target for reaching a predicted atmospheric entry location, wherein the predicted atmospheric entry location corresponds to a point at which the spacecraft interacts with an atmosphere in order to reach the target landing site, (ii) determining, by the computing device, a back-propagated orbit state estimate of the spacecraft, wherein the back-propagated orbit state estimate of the spacecraft corresponds to an estimated location and path-angle of the spacecraft from the predicted atmospheric entry location based on the range target and the velocity target, (iii) comparing, by the computing device, the back-propagated orbit state estimate to a known orbit state of the spacecraft to determine that the back-propagated orbit state estimate has converged with the known orbit state, and (iv) calculating, by the computing device, based on determining that the back-propagated orbit state estimate has converged with the known orbit state, (a) an estimated time of ignition for a propulsion system of the spacecraft and (b) an estimated burn velocity vector of the propulsion system using the range target and the velocity target; (i) determining, by the computing device, a range target and a velocity target for reaching a predicted atmospheric entry location, wherein the predicted atmospheric entry location corresponds to a point at which the spacecraft interacts with an atmosphere in order to reach the target landing site, (ii) determining, by the computing device, a back-propagated orbit state estimate of the spacecraft, wherein the back-propagated orbit state estimate of the spacecraft corresponds to an estimated location and path-angle of the spacecraft from the predicted atmospheric entry location based on the range target and the velocity target, (iii) comparing, by the computing device, the back-propagated orbit state estimate to a known orbit state of the spacecraft to determine that the back-propagated orbit state estimate has converged with the known orbit state, and (iv) calculating, by the computing device, based on determining that the back-propagated orbit state estimate has converged with the known orbit state, (a) an estimated time of ignition for a propulsion system of the spacecraft and (b) an estimated burn velocity vector of the propulsion system using the range target and the velocity target; and at or prior to an end time of the predetermined time window, performing a burn pulse by the propulsion system in accordance with the estimated time of ignition and the estimated burn velocity vector.