Systems and methods for autonomous deorbiting of a spacecraft

A system for deorbiting a spacecraft comprising:

• a spacecraft comprising:
• a computing device having a processor and memory storing instructions executable by the processor to:
• select a target landing site for deorbiting the spacecraft;
• during a predetermined time window associated with deorbiting the spacecraft:
• (i) determine 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) determine, 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) compare 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) calculate 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
• the propulsion system configured to, at or prior to an end time of the predetermined time window, perform a burn pulse in accordance with the estimated time of ignition and the estimated burn velocity vector.

A non-transitory computer readable medium having stored thereon instructions, that when executed by one or more processors of a computing device, cause the computing device to perform functions comprising:

• selecting a target landing site for deorbiting a spacecraft;
• during a predetermined time window associated with deorbiting the spacecraft:
• (i) determining 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 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 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 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.