Hybrid rocket motor with annular, concentric solid fuel elements
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02K-009/28
출원번호
US-0764677
(2007-06-18)
등록번호
US-8539753
(2013-09-24)
발명자
/ 주소
Macklin, Frank
출원인 / 주소
SpaceDev, Inc.
대리인 / 주소
Hernandez, Fred C.
인용정보
피인용 횟수 :
0인용 특허 :
16
초록▼
A hybrid rocket motor includes a supply of oxidizer, a first solid fuel element positioned around the supply of oxidizer, a second solid fuel element positioned concentrically around the first solid fuel element, and a combustion port positioned between the first and second solid fuel elements. The
A hybrid rocket motor includes a supply of oxidizer, a first solid fuel element positioned around the supply of oxidizer, a second solid fuel element positioned concentrically around the first solid fuel element, and a combustion port positioned between the first and second solid fuel elements. The oxidizer interacts with the first and second solid fuel elements within the combustion port to produce a combustion product. A nozzle is in communication with the combustion port for combustion discharge of the combustion product.
대표청구항▼
1. A hybrid rocket motor, comprising: an oxidizer tank containing a supply of oxidizer;a first, annular solid fuel element positioned around the supply of oxidizer;a second solid fuel element positioned concentrically around the first solid fuel element;a combustion port positioned between the first
1. A hybrid rocket motor, comprising: an oxidizer tank containing a supply of oxidizer;a first, annular solid fuel element positioned around the supply of oxidizer;a second solid fuel element positioned concentrically around the first solid fuel element;a combustion port positioned between the first and second solid fuel elements, wherein the oxidizer interacts with the first and second solid fuel elements within the combustion port to produce a combustion product; anda nozzle in communication with the combustion port for combustion discharge of the combustion product;wherein a generally enclosed space is positioned below the oxidizer tank and below the first solid fuel element, such that the widest diameter of the generally enclosed space is substantially equal to the outer diameter of the first, annular solid fuel element, wherein the space communicates with a throat of the nozzle via at least one passageway therebetween and wherein the at least one passageway can be used to facilitate liquid injection vector thrust control of the nozzle. 2. A hybrid rocket motor as in claim 1, wherein the second solid fuel element is annular. 3. A hybrid rocket motor as in claim 1, wherein the oxidizer comprises Nitrous Oxide. 4. A hybrid rocket motor as in claim 1, wherein the first and second solid fuel elements comprise polymethyl methacrylate, high density polyethylene, or hydroxyl terminated polybutadiene. 5. A hybrid rocket motor as in claim 1, wherein the first and second solid fuel elements are contained within a solid fuel tank that surrounds the supply of oxidizer. 6. A hybrid rocket motor as in claim 5, further comprising at least one injector adapted to inject oxidizer from the supply of oxidizer into the solid fuel tank. 7. A hybrid rocket motor as in claim 5, further comprising a plurality of injectors spaced around a circumference of the solid fuel tank. 8. A hybrid rocket motor as in claim 1, wherein the hybrid rocket motor comprises an aerospike. 9. A hybrid rocket motor, comprising: an oxidizer tank containing an oxidizer; a main casing surrounding the oxidizer tank; at least one injector adapted to inject oxidizer from the oxidizer tank into the main casing; wherein the main casing includes: a first, annular solid fuel element;a second, annular solid fuel element positioned concentrically around the first, annular solid fuel element;a combustion port positioned between the first and second annular solid fuel elements wherein the oxidizer interacts with the first and second solid fuel elements within the combustion port to produce a combustion product; anda nozzle in communication with the combustion port for combustion discharge of the combustion product;wherein a generally enclosed space is positioned below the oxidizer tank and below the first, annular solid fuel element, such that the widest diameter of the generally enclosed space is substantially equal to the outer diameter of the first, annular solid fuel element, wherein the space communicates with a throat of the nozzle via at least one passageway therebetween and wherein the at least one passageway can be used to facilitate liquid injection vector thrust control of the nozzle. 10. A hybrid rocket motor as in claim 9, wherein the hybrid rocket motor comprises an aerospike. 11. A hybrid rocket motor as in claim 9, wherein the at least one injector comprises a plurality of injectors spaced around a circumference of the main casing. 12. A hybrid rocket motor as in claim 9, wherein the oxidizer comprises Nitrous Oxide. 13. A hybrid rocket motor as in claim 9, wherein the first and second solid fuel elements comprise polymethyl methacrylate, high density polyethylene, or hydroxyl terminated polybutadiene. 14. A hybrid rocket motor as in claim 9, further comprising a third solid fuel element concentrically surrounding the second solid fuel element.
Schumacher John B. ; Kosky John P., Encapsulated propellant grain composition, method of preparation, article fabricated therefrom and method of fabricatio.
Bradford Michael D. (4896 Kenneth Ave. Santa Maria CA 93455) Kniffen ; Jr. Roy J. (934 Sharon La. ; #1 Ventura CA 93001) McKinney Bevin C. (252 N. Crimea Ventura CA 93004), Hybrid rocket combustion enhancement.
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