Method and apparatus for a substantially coaxial injector element

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.

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.

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.

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.

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.

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.

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.

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.