Granted Patent B2

US 7,124,574 · App. 10/309,833 · Granted Oct 24, 2006

Abstract

Claims (131)

1. A space craft engine, comprising:

a propellant source;

a combustion chamber;

a face plate having a propellant side and a combustion side;

a propellant injector formed in said faceplate, including:

a first propellant annulus extending through said faceplate;

a second propellant annulus defined by a counterbore in said faceplate arranged substantially coaxial with said first propellant annulus; and

wherein said coaxial arrangement of said second propellant annulus relative to said first propellant annulus allows a propellant from said propellant source to be injected into said combustion chamber and thoroughly intermixed prior to being combusted;

an inter-propellant plate to divide a first propellant and a second propellant in said propellant source;

said first propellant being injected into said combustion chamber through said first propellant annulus and said second propellant being injected into said combustion chamber through said second propellant annulus; and

wherein said first propellant includes an oxidizer and said second propellant includes a fuel.

2. The space craft engine of claim 1 , wherein:

said faceplate includes a throughbore through which said elongated tube extends; and

said second propellant annulus is formed from said throughbore.

3. The space craft engine of claim 2 , further comprising:

a passageway defined in said faceplate so that a volume of propellant is able to flow to said combustion chamber through said second propellant annulus.

4. The space craft engine of claim 1 , wherein a radius of said second propellant annulus is adapted to control the flow of a volume of propellant from said propellant source to said combustion chamber.

5. A space craft engine comprising:

an injector to inject a propellant into a combustion chamber of a combustion system, including:

a faceplate having a first side and a second side and a bore extending between said first side and said second side;

a propellant tube defining a first propellant annulus adapted to engage a portion of said bore;

said faceplate defining a second propellant annulus substantially concentric with said propellant tube; and

wherein said propellant tube and said second propellant annulus define a path to deliver a volume of propellant to a combustion chamber.

6. The space craft engine of claim 5 , further comprising:

a passageway defined though said faceplate extending from said first side of said faceplate to said second propellant annulus; and

wherein said passageway is adapted to allow a selected flow rate of the propellant to said second propellant annulus.

7. The space craft engine of claim 5 ,

wherein a radius of said second annulus defines a flow rate of said propellant through said second propellant annulus; and

wherein a ratio of a radius of said propellant tube and said radius of said second propellant annulus substantially defines a mix rate of said propellant.

8. For a rocket engine including a faceplate, having a propellant side and a combustion side, and an injector formed in the faceplate including a throughbore extending between the propellant side and the combustion side, the injector comprising:

a propellant tube adapted to extend through a portion of said throughbore;

a propellant annulus defined in the faceplate around said throughbore wherein a center of said throughbore defines a center of said propellant annulus;

a passageway extending from the propellant side of the faceplate to said propellant annulus; and

wherein a propellant is adapted to flow from said propellant side to said combustion side though the injector;

wherein said passageway is adapted to allow a selected flow rate of the propellant to said propellant annulus.

9. The injector of claim 8 , comprising:

wherein said passageway includes a plurality of said passageways;

wherein said plurality of passageways are disposed around said propellant tube at intervals of about 90°; and

wherein each of said plurality of passageways extends to said propellant annulus.

10. The injector of claim 8 , wherein said propellant tube defines a second propellant annulus.

11. The injector of claim 10 , wherein:

a radius of said propellant annulus defines a flow rate of said propellant through said propellant annulus; and

a ratio of a radius of said propellant tube and said radius of said propellant annulus substantially defines a mix rate of said propellant.

12. The injector of claim 8 , wherein an external dimension of said propellant tube is substantially equal to an internal dimension of said throughbore.

13. A method of forming an injector for a combustion powered engine comprising:

providing a faceplate having a first side and a second side;

forming a throughbore in said faceplate that extends between said first side and said second side;

enlarging a portion of said throughbore thereby forming a first propellant annulus;

forming a second propellant annulus that is coaxial with said first propellant annulus; and

wherein a propellant is able to flow though said first propellant annulus and said second propellant annulus.

14. The method of claim 13 , wherein providing a faceplate includes:

forming a material into a substantially planar surface; and

disposing an inter-propellant plate a distance from said first side.

15. The method of claim 13 , wherein forming a throughbore includes:

selecting a propellant tube outer diameter;

forming said throughbore to have a tolerance of less than 0.02 mm relative to the propellant tube outer diameter; and

wherein said tolerance substantially eliminates a flow of a propellant between a propellant tube and said faceplate.

16. The method of claim 13 , wherein forming a first propellant annulus includes:

providing a drill bit including a pilot portion and a cutting portion, wherein said cutting portion includes a cutting radius larger than said pilot portion;

disposing said pilot portion in said throughbore;

guiding said cutting portion into said faceplate with said pilot portion; and

wherein said cutting radius includes a radius selected in relation to a desired flow rate of a propellant.

17. The method of claim 13 , wherein forming a second propellant annulus includes:

disposing a propellant tube in said throughbore such that said propellant tube extends at least between said first side and said second side.

18. The method of claim 13 , further comprising:

forming a passageway through said faceplate between said first side of said faceplate and said first propellant annulus; and

wherein said passageway allows a selected flow rate of a propellant to said first propellant annulus.

19. A method of forming an injector for a combustion powered engine comprising:

providing a faceplate having a first side and a second side;

forming a throughbore in said faceplate to extend between said first side and said second side;

forming a first propellant annulus by enlarging a portion of said throughbore by guiding a cutting bit into said faceplate with said throughbore;

disposing a propellant tube in said throughbore to form a second propellant annulus; and

forming a passageway in said faceplate between said first side and said first propellant annulus.

20. The method of claim 19 , wherein forming a first propellant annulus includes:

providing the cutting bit with a pilot portion and a cutting portion, wherein said cutting portion includes a cutting radius larger than said pilot portion;

disposing said pilot portion in said throughbore; and

guiding said cutting portion into said faceplate with said pilot portion.

21. The method of claim 19 , wherein forming a throughbore includes:

selecting a propellant tube outer diameter;

forming said throughbore to have generally tight tolerances to the propellant tube outer diameter; and

wherein generally tight tolerances substantially eliminates a flow of a propellant between a propellant tube and said faceplate.

22. A spacecraft comprising:

a payload compartment;

a propellant source;

an engine including:

a combustion chamber;

a propellant chamber;

a faceplate disposed between said combustion chamber and said propellant chamber;

a propellant injector formed in said faceplate, including:

a first propellant annulus extending through said faceplate;

a second propellant annulus defined by a counterbore in said faceplate arranged substantially coaxial with said first propellant annulus; and

wherein said substantial coaxial arrangement of said second propellant annulus relative to said first propellant annulus allows a propellant from said propellant source to be injected into said combustion chamber and thoroughly intermixed prior to being combusted.

23. An apparatus to allow for substantial mixing of a first fluid with a second fluid into a common chamber, comprising:

an injector to inject the first fluid or the second fluid into the common chamber, including:

a faceplate having a first side and a second side and a bore extending between said first side and said second side;

an injector tube cannula defining a first fluid annulus adapted to engage a portion of said bore;

said faceplate defining a second injector annulus substantially concentric with said injection tube cannula; and

wherein said injector tube and said second injector annulus define a path to deliver a volume of the first fluid or the second fluid to the common chamber.

24. The apparatus of claim 23 , further comprising:

a passageway defined through said faceplate extending from said first side of said faceplate to said second injector annulus;

wherein said passageway is adapted to allow a selected flow rate of the first fluid or the second fluid to said second injector annulus.

25. The apparatus of claim 23 ,

wherein a radius of said second injector annulus defines a flow rate of the first fluid or the second fluid through said second injector annulus; and

wherein a ratio of a radius of said injector tube and said radius of said second injector annulus substantially defines a mix rate of the first fluid and the second fluid.

26. An engine, comprising:

a first propellant source operable to provide a source of a first propellant;

a second propellant source operable to provide a source of a second propellant;

an inter-propellant plate separating said first propellant source and said second propellant source;

a combustion chamber operable to contain a portion of a combustion of the first propellant and the second propellant;

a face plate having a propellant side and a combustion side;

a propellant injector formed in said faceplate, including:

a propellant tube extending through said faceplate and said interpropellant plate;

a propellant bore defined by a bore in said faceplate arranged near said propellant tube; and

wherein said arrangement of said propellant bore relative to said propellant tube allows the first propellant and the second propellant from said first propellant source and said second propellant source to be injected into said combustion chamber; and

a propellant passage operable to transport at least one of the first propellant or the second propellant from at least one of the first propellant source or the second propellant source to said propellant bore.

27. The engine of claim 26 , wherein said propellant tube extends through said inter-propellant plate to allow passage of the first propellant to the combustion chamber and said propellant passage extends through a portion of said face plate to interconnect the second propellant source and said propellant bore.

28. The engine of claim 26 , wherein each of said propellant tube, said propellant bore, or said propellant passage defines an annulus.

29. The engine of claim 26 , wherein a size of at least on of said propellant tube, said propellant bore, said propellant passage, or combinations thereof define a flow rate of at least one of the first propellant or the second propellant from at least one of said first propellant source or said second propellant source to said combustion chamber.

30. A space craft engine, comprising:

a propellant source;

a combustion chamber;

a face plate having a propellant side and a combustion side;

a propellant injector formed in said faceplate, including:

a first propellant passage member extending through said faceplate;

a counterbore formed in said faceplate and defining a second propellant passage relative to said first propellant passage member; and

an inter-propellant plate to divide a first propellant and a second propellant in said propellant source;

wherein said coaxial arrangement of said second propellant annulus relative to said first propellant annulus allows a propellant from said propellant source to be injected into said combustion chamber and thoroughly intermixed prior to being combusted;

wherein said first propellant being injected into said combustion chamber through said first propellant annulus and said second propellant being injected into said combustion chamber through said second propellant annulus;

wherein said first propellant includes an oxidizer and said second propellant includes a fuel;

wherein said first annulus is defined by an elongated tube disposed substantially in a center of said second annulus; and

wherein said elongated tube operably engages said inter-propellant plate.

Assignments (12)