Spin-stabilized orbital rocket guidance
| DWPI Title: Spin-stabilized rocket, has hardware logic device for causing first-stage engine to initiate burn of first stage engine at computed ignition time when rocket has computed orientation, where burn causes rocket to follow path to trajectory |
| Abstract: Technologies for guidance of a spin-stabilized orbital rocket are described herein. The spin-stabilized rocket includes a guidance controller. The guidance controller computes parameters of a burn of a second-stage engine of the rocket to reach a desired nominal orbit subsequent to burnout of the first stage of the rocket. The guidance controller computes the burn parameters of the second-stage engine based upon one or more desired orbit parameters and a current position and velocity of the rocket. The computation of the burn parameters is based upon a simulated point-mass model of the motion of the rocket. The guidance controller then controls the rocket to initiate a second-stage burn having the computed burn parameters. |
| Use: Spin-stabilized rocket i.e. three-stage rocket, for maneuvering to a nominal orbit trajectory without thrust vector control. |
| Advantage: The spin-stabilized rocket is configured to maneuver to a nominal orbit trajectory without thrust vector control. The thrust vectoring system reduces size, weight, complexity, and cost of a rocket. The rocket obtains burn parameters of the first stage simultaneously with the second stage and allows a guidance controller to compute second stage burn parameters that place the rocket on a trajectory that allows to take the rocket to the nominal orbit. The guidance controller computes burn parameters for succeeding stages after the stage completes burn. |
| Novelty: The rocket (100) has a hardware logic device for controlling operation of a first-stage engine and a second-stage engine i.e. solid-fuel rocket engine. The hardware logic device computes an orientation of the spin-stabilized rocket for a burn of the second-stage engine and an ignition time of the burn of the second-stage engine based upon a desired orbit parameter, a current velocity of the rocket, and a current position of the rocket after completion of a burn of the first-stage engine. The hardware logic device causes the second-stage engine to initiate burn of the second-stage engine at computed ignition time when the rocket has computed orientation. A reaction control system (RCS) comprises a thruster. |
| Filed: 7/24/2019 |
| Application Number: US16521001A |
| Tech ID: SD 13330.1 |
| This invention was made with Government support under Contract No. DE-NA0003525 awarded by the United States Department of Energy/National Nuclear Security Administration. The Government has certain rights in the invention. |
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