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OwnerErich Kirner, Helmut Dederrainventors
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A method of operating a liquid fuel rocket engine having at least first and second propellant component pumps, driven by a gas turbine which is connected to discharge its turbine propulsion gases through a secondary thrust nozzle, includes a main engine having a combustion chamber with a thrust nozzle, an initial expansion portion with an expansion corresponding to ground pressure and a final expansion portion connected to the initial expansion portion and terminating in a thrust discharge operating substantially at vacuum and with means for circulating at least one propellant component through cooling channels of the walls of both expansion portions, comprises, directing a first component into the combustion chamber, and directing at least a first portion of a second propellant component into the walls of the initial expansion portion and then into the combustion chamber and directing a second portion of the second propellant component into the walls of the final expansion portion so that they are heated therein and, thereafter, directing the heated propellant component gases into the turbine. The liquid fuel rocket engine for operation in outer space, comprises, wall means defining a combustion chamber having an injection head and a thrust nozzle part with a narrow neck portion, an initial expanding portion and a final expanding portion which terminates in a thrust discharge.
A liquid fuel rocket engine for operation in outer space, comprising, wall means defining a combustion chamber having an injection head and with a thrust nozzle part having a narrow neck portion, an initial expanding portion having an expansion of substantially ground pressure and a vacuum expanding portion connected to said initial expanding portion and terminating in a thrust discharge, each of said combustion chamber, said initial expansion portion and said final expanding portion having walls with cooling channels extending therethrough for the circulation of a propellant component therethrough, a propellant component pump gas turbine, first and second propellant component pumps connected to said gas turbine for operation thereby, a first propellant component supply tank connected to said first propellant component pump, a second propellant component supply tank connected to said second propellant component pump, a secondary flow nozzle connected to said turbine for the discharge of the turbine exhaust gases, a distributor ring connected to the cooling channels of said initial expansion portion and said final expansion portion for circulating a first portion of said second propellant component to the cooling channels of said initial expansion portion, and a second portion of said second propellant component to the cooling channels of said second expansion portion, a first collecting ring connected to said final expansion portion and disposed adjacent said discharge, a second circulating line connected to said second propellant component pump and said distributor ring for supplying a second propellant component to said distributor ring, a second collecting ring connected to said cooling channels of said initial expanding portion and disposed adjacent said head, and a first connection line from said collecting ring to said turbine for directing a portion of said second propellant component gases to said turbine after they have been heated in the cooling channels of said final expansion portion.
A method of operating a liquid fuel rocket engine having at least a first and a second propellant component pump driven by a gas turbine, which is connected to discharge its turbine propulsion gases through a secondary thrust nozzle which includes a main combustion chamber with a thrust nozzle and an initial expansion portion with an expansion corresponding to ground pressure, and a final expansion portion connected to the initial expansion portion which terminates in a thrust discharge operating at substantially vacuum and with means for circulating at least one portion of a first propellant component through the cooling channels of the walls, the initial expansion portion and the combustion chamber, and at least one second portion through the channels of the walls of the final expansion portion, comprising, directing the first propellant into the combustion chamber, directing a first portion of the second propellant component into the initial expansion portion and the combustion chamber cooling channel and directing a second portion into the walls of the final expansion portion so that they are heated therein and directing the heated portion of the second propellant component into the turbine to operate the turbine.
Dependencies are inferred from the claim text.