Closed Loop Guidance Algorithm For Low Thrust Ballistic Re-entry Vehicle
- Paper number
IAC-07-C1.6.08
- Author
Mr. Rajeev U.P., Indian Space Research Organisation (ISRO), VSSC, India
- Coauthor
Mr. Sheela D.S., Indian Space Research Organisation (ISRO), VSSC, India
- Coauthor
Mr. U. Janardhana Naik, Indian Space Research Organisation (ISRO), VSSC, India
- Coauthor
Dr. Bijan B Das, Indian Space Research Organisation (ISRO), VSSC, India
- Coauthor
Dr. S. Dasgupta, India
- Year
2007
- Abstract
A closed loop guidance algorithm to guide a ballistic re-entry vehicle from orbit to atmospheric entry is presented. The mission begins with a de-boost phase where Eight Reaction Control System (RCS) thrusters are used to reduce the velocity of the vehicle. Since the thrust level of the RCS system is very low; non-linear, pseudo forces like Coriolis force and Euler force are dominant in the vehicle dynamics. Hence conventional launch vehicle guidance algorithms that are developed for high thrust propulsion systems cannot be applied directly for this mission. Augmented Flat Earth (AFE) Guidance developed in this paper is a uniform gravity based algorithm in which the pseudo forces are accounted. After de-boost, the re-entry trajectory has a long coast till atmospheric entry and hence Earth oblateness effects are modeled in the algorithm using an on-line re-targeting technique. For guidance development, the entry trajectory of the vehicle is divided into:
(1) De-boost phase, where the low thrust propulsion system transfers the vehicle from the initial orbit to the descending leg of an entry ellipse.
(2) Coast phase, where the vehicle coasts to atmospheric entry point ensuring zero angle of attack at entry so that during atmospheric flight, the vehicle is aerodynamically stabilized.
(3) Atmospheric phase where the vehicle follows a ballistic trajectory till touchdown.
The novel features of the guidance system described in this paper are the following.
(1) De-boost phase guidance algorithm uses models of Euler force, Coriolis force and the low thrust propulsion force.
(2) The algorithm steers the vehicle to an inertially fixed descending orbit using an on-line target prediction technique.
(3) Though the forces modeled are non-linear functions of states, a closed form solution is derived to ensure reliable on-board implementation.
(4) During the de-boost phase and coast phase, guidance ensures the GPS antenna of the vehicle points to the outer space.
(5) The algorithm uses an on-line retargeting to compensate the oblateness effects of Earth’s gravitational field.
(6) The algorithm can be used for general, propulsive orbit transfer missions.
Results of 6D simulations ensure the desired entry conditions even under large variations in propulsion parameters and initial conditions. The resulting down-range and cross-range errors is within 1 km for near nominal performances. For thrust perturbations up to 30 percent, the down-range and cross-range errors are within 20km. This establishes the robustness of the algorithm. Pre-flight simulation results and flight experience are discussed. Post flight analysis shows that all the mission objectives are satisfied.
- Abstract document
- Manuscript document
IAC-07-C1.6.08.pdf (🔒 authorized access only).
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