Probability and frequency of orbital encounters
Systems, methods, devices, and non-transitory media of the various embodiments provide for a volumetric approach to determining orbital encounters that may determine the number of encounters over a specified length of time. Such information may be used to determine how often during an orbit or period of time an object might trigger a conjunction warning for a neighboring satellite. The various embodiments may be used as a planning and characterization tool to estimate satellite encounter rates for a prospective orbit regime and may provide an efficient, in-line approach to assess the number of encounters occurring within a user-specified span of time.
1. A method for determining an encounter probability between satellites, comprising:
receiving initial orbital elements of a first satellite, the initial orbital elements of the first satellite comprising an orbit semi-major axis, an eccentricity, an inclination, an ascending node, an argument of perigee, and a mean anomaly;
calculating a first satellite's orbit according to the initial orbital elements of the first satellite;
receiving orbital elements of a second satellite;
calculating an orbit of a second satellite according to orbital elements of the second satellite;
receiving characteristics of a volumetric shape that encloses the second satellite;
rotating the volumetric shape along the second satellite's orbit to create a toroid;
determining each penetration of the toroid by the first satellite's orbit, wherein each penetration comprises a nodal crossing;
determining a length of each nodal crossing;
calculating an instantaneous encounter probability value based on a total length of the nodal crossings for a single revolution about the orbit of the second satellite; and
using the instantaneous encounter probability value to adjust the first satellite's orbit.
2. The method of claim 1 , wherein the initial orbital elements of the first satellite create a closed orbital path.
3. The method of claim 1 , wherein the mean anomaly is a true anomaly for the first satellite.
4. The method of claim 1 , wherein the mean anomaly is chosen from a range to force an initial encounter.
5. The method of claim 1 , wherein the initial orbital elements of the first satellite further comprise J2 terms, oblate central body effects, drag, solar radiation pressure (SRP), and 3rd-body gravity influences.
6. The method of claim 1 , wherein the volumetric shape comprises an irregular enclosed volume.
7. The method of claim 1 , wherein the volumetric shape further comprises a geometric volume.
8. The method of claim 7 , wherein the volumetric shape is an ellipsoid.
9. The method of claim 7 , wherein the volumetric shape is a sphere.
10. The method of claim 1 , further comprising:
determining by an apogee-perigee filter whether the first satellite's orbit comes within a range of the second satellite's orbit before determining each penetration of the toroid by the first satellite's orbit; and
determining each penetration of the toroid by the first satellite's orbit in response to determining by the apogee-perigee filter that the first satellite's orbit comes within the range of the first satellite's orbit.
11. A computing device, comprising:
a processor, wherein the processor is configured with processor-executable instructions to perform operations comprising:
receiving initial orbital elements of a first satellite, the initial orbital elements of the first satellite comprising an orbit semi-major axis, an eccentricity, an inclination, an ascending node, an argument of perigee, and a mean anomaly;
calculating a first satellite's orbit according to the initial orbital elements of the first satellite;
receiving orbital elements of a second satellite;
calculating an orbit of a second satellite according to orbital elements of the second satellite;
receiving characteristics of a volumetric shape that encloses the second satellite;
rotating the volumetric shape along the second satellite's orbit to create a toroid;
determining each penetration of the toroid by the first satellite's orbit, wherein each penetration comprises a nodal crossing;
determining a length of each nodal crossing;
calculating an instantaneous encounter probability value based on a total length of the nodal crossings for a single revolution about the second satellite's orbit; and
using the instantaneous encounter probability value to adjust the first satellite's orbit.
12. The computing device of claim 11 , wherein the initial orbital elements of the first satellite create a closed orbital path.
13. The computing device of claim 11 , wherein the mean anomaly is a true anomaly for the first satellite.
14. The computing device of claim 11 , wherein the mean anomaly is chosen from a range to force an initial encounter.
15. The computing device of claim 11 , wherein the initial orbital elements of the first satellite further comprise J2 terms, oblate central body effects, drag, solar radiation pressure (SRP), and 3rd-body gravity influences.
16. The computing device of claim 11 , wherein the volumetric shape comprises an irregular enclosed volume.
17. The computing device of claim 11 , wherein the volumetric shape further comprises a geometric volume.
18. The computing device of claim 17 , wherein the volumetric shape is an ellipsoid.
19. The computing device of claim 17 , wherein the volumetric shape is a sphere.
20. The computing device of claim 11 , wherein the processor is configured with processor-executable instructions to perform operations further comprising:
determining by an apogee-perigee filter whether the first satellite's orbit comes within a range of the second satellite's orbit before determining each penetration of the toroid by the first satellite's orbit; and
determining each penetration of the toroid by the first satellite's orbit in response to determining by the apogee-perigee filter that the first satellite's orbit comes within the range of the second satellite's orbit.