Systems and methods for autonomous deorbiting of a spacecraft

Patent metadata

Field	Value
Publication number	US-11465782-B2
Application number	US-201916553206-A
Country	US
Kind code	B2
Filing date	Aug 28, 2019
Priority date	Aug 28, 2019
Publication date	Oct 11, 2022
Grant date	Oct 11, 2022

Abstract

In an example, a method for deorbiting a spacecraft is described. The method includes selecting a target landing site for deorbiting the spacecraft. The method includes determining a range target and a velocity target for reaching a predicted atmospheric entry location. The method includes determining a back-propagated orbit state estimate of the spacecraft. The method includes 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. The method includes 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. The method includes performing a burn pulse by the propulsion system.

First claim

What is claimed is:

1. A method of deorbiting a spacecraft, comprising:
receiving, by way of a user interface of a computing device, a selection of a predetermined time window for deorbiting the spacecraft;
selecting, by the computing device, a target landing site based on a known orbit state of the spacecraft and the predetermined time window for deorbiting the spacecraft;
during the predetermined time window:
(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 the 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.

2. The method of claim 1 , wherein selecting the target landing site for deorbiting the spacecraft comprises determining a priority level associated with each of a plurality of ground locations based at least on a location type associated with each location; and
selecting a landing site having a highest priority level.

3. The method of claim 2 , wherein the spacecraft comprises a database of ground locations, and wherein determining the priority level associated with each of the plurality of ground locations comprises retrieving each priority level from the database of ground locations.

4. The method of claim 1 , wherein selecting the target landing site for deorbiting the spacecraft comprises:
determining, based on a map of target ground locations, that no designated landing facilities on the map are reachable based on the known orbit state of the spacecraft and the predetermined time window for deorbiting the spacecraft; and
responsive to determining that no designated landing facilities are reachable, selecting an emergency landing location that is reachable based on the known orbit state of the spacecraft and the predetermined time window for deorbiting the spacecraft, wherein the emergency landing location corresponds to a water zone polygon defining an area on a water surface.

5. The method of claim 4 , wherein selecting the emergency landing location comprises:
determining that an orbit path of the spacecraft crosses a boundary of the water zone polygon; and
responsive to determining that the orbit path of the spacecraft crosses the boundary of the water zone polygon, selecting the emergency landing location.

6. The method of claim 1 , further comprising:
determining an expected arrival time for the spacecraft at each of a plurality of ground locations, wherein selecting the target landing site for deorbiting the spacecraft from the plurality of ground locations comprises:
determining a landing site score for each ground location based on each corresponding expected arrival time and based on each corresponding priority level; and
selecting, as the target landing site for deorbiting the spacecraft, a particular ground location having a highest landing site score from the plurality of ground locations.

7. The method of claim 1 , further comprising:
receiving, from one or more sensors of the spacecraft, sensor data indicative of at least a position and velocity of the spacecraft; and
determining the known orbit state of the spacecraft based on the sensor data.

8. The method of claim 1 , further comprising determining a type of the propulsion system, wherein calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system comprises calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system based on the type of the propulsion system.

9. The method of claim 1 , wherein calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system comprises calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system based on a desired path-angle of the spacecraft and a desired velocity of the spacecraft relative to the atmosphere at the predicted atmospheric entry location.

10. The method of claim 1 , further comprising, during the predetermined time window associated with deorbiting the spacecraft:
determining a set of range targets and velocity targets having back-propagated orbit state estimates that converge with corresponding known orbit states of the spacecraft; and
selecting an optimal range target and velocity target from the set of range targets and velocity targets that has a back-propagated orbit state estimate that most closely converges with a corresponding known orbit state,
wherein the range target and the velocity target correspond to the optimal range target and velocity target from the set of range targets and velocity targets, and
wherein calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system using the range target and the velocity target, comprises calculating the estimated time of ignition for the propulsion system of the spacecraft and the estimated burn velocity vector of the propulsion system using the optimal range target and velocity target from the set of range targets and velocity targets.

11. The method of claim 1 , further comprising:
determining a forward-propagated atmospheric interface, wherein the forward-propagated atmospheric interface corresponds to expected force characteristics to be experienced by the spacecraft when entering the atmosphere based on the estimated time of ignition, the estimated burn velocity vector, and the known orbit state of the spacecraft;
determining an estimated range error and an estimated path-angle error of the spacecraft at the forward-propagated atmospheric interface based on the expected force characteristics;
determining that the estimate range error converged within a threshold range error;
determining that the estimated path-angle error has converged within a threshold path-angle error; and
calculating an actual time of ignition and an actual burn velocity vector based on (i) determining that the estimate range error converged within a threshold range error and (ii) determining that the estimated path-angle error has converged within a threshold path-angle error,
wherein performing the burn pulse by the propulsion system in accordance with the estimated time of ignition and the estimated burn velocity vector comprises performing the burn pulse by the propulsion system at the

Assignees

• Boeing Co

Inventors

• Ealy John P
• Phail Brian A
• Poladian Dave

Classifications

• B64G1/24 Primary

  Guiding or controlling apparatus, e.g. for attitude control · CPC title

• G01C23/00

  Combined instruments indicating more than one navigational value, e.g. for aircraft; Combined measuring devices for measuring two or more variables of movement, e.g. distance, speed or acceleration · CPC title

• B64G1/36

  using sensors, e.g. sun-sensors, horizon sensors · CPC title

• B64G1/242 Primary

  Orbits and trajectories · CPC title

• B64G2001/247

  Operations & Transport · mapped topic