Dual-vortical-flow hybrid rocket engine
The present invention discloses a dual-vortical-flow hybrid rocket engine, including a main body and a nozzle communicating with an end of the main body. The main body includes a plurality of disk-like combustion chambers arranged longitudinally, and a central combustion chamber formed along the axial portion and communicating the disk-like combustion chambers. Each of the disk-like combustion chambers is provided with a plurality of oxidizer injection nozzles at its inner circumference surface. Inside the disk-like combustion chambers, the oxidizer is injected in nearly the tangent directions of the circumference, and the injection directions are opposite for the neighboring disk-like combustion chambers, which creates vortical flows with opposite rotating directions so as to increase the total residence time of the combustion reactions of the oxidizer and the solid-state fuel in the disk-like combustion chambers of the present invention.
1. A dual-vortical-flow rocket engine comprising:
a main body having:
a plurality of disk-like combustion chambers arranged longitudinally;
a central combustion chamber communicating with an axial portion of each disk-like combustion chamber; and
a plurality of oxidizer injection nozzles disposed at an inner circumference surface of each disk-like combustion chamber, wherein the plurality of oxidizer injection nozzles of one of the plurality of disk-like combustion chambers face toward a clockwise direction of the inner circumference surface of the one the plurality of disk-like combustion chambers, and the plurality of oxidizer injection nozzles of another one of the plurality of disk-like combustion chambers face toward a counterclockwise direction of the inner circumference surface of the another one the plurality of disk-like combustion chambers, such that a rotating direction of the oxidizer injected into the one of the plurality of disk-like combustion chambers is clockwise, and a rotating direction of the oxidizer injected into the another one of the plurality of disk-like combustion chambers is counterclockwise, wherein the one the plurality of disk-like combustion chambers and the another one of the plurality of disk-like combustion chambers are two immediately adjacent disk-like combustion chambers; and
a nozzle connected at an end of the main body.
2. The dual-vortical-flow rocket engine as claim 1 , wherein, each oxidizer injection nozzle and a corresponding tangent line of the circumference surface of each disk-like combustion chamber form a 10˜20 degree angle such that an oxidizer injects into each disk-like combustion chamber with the 10˜20 degree angles measured from the corresponding tangent line of the circumference surface of each disk-like combustion chamber.
3. The dual-vortical-flow rocket engine as claim 2 , wherein, the number of the disk-like combustion chambers is even.
4. The dual-vortical-flow rocket engine as claim 3 , wherein, the number of the disk-like combustion chambers is two.
5. The dual-vortical-flow rocket engine as claim 1 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is even and the oxidizer injection nozzles disposes symmetrically.
6. The dual-vortical-flow rocket engine as claim 2 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is even and the oxidizer injection nozzles disposes symmetrically.
7. The dual-vortical-flow rocket engine as claim 3 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is even and the oxidizer injection nozzles disposes symmetrically.
8. The dual-vortical-flow rocket engine as claim 4 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is even and the oxidizer injection nozzles disposes symmetrically.
9. The dual-vortical-flow rocket engine as claim 5 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is at least four.
10. The dual-vortical-flow rocket engine as claim 6 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is at least four.
11. The dual-vortical-flow rocket engine as claim 7 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is at least four.
12. The dual-vortical-flow rocket engine as claim 8 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is at least four.
13. The dual-vortical-flow rocket engine as claim 9 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is four, six, or eight.
14. The dual-vortical-flow rocket engine as claim 10 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is four, six, or eight.
15. The dual-vortical-flow rocket engine as claim 11 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is four, six, or eight.
16. The dual-vortical-flow rocket engine as claim 12 , wherein, the number of the oxidizer injection nozzles disposed at the inner circumference surface of each disk-like combustion chamber is four, six, or eight.