Abstract: A liquid-fueled rocket engine assembly comprises a combustor assembly, a combustor jacket for cooling the combustor assembly using pressurized liquid fuel exiting a fuel pump, a turbine for expanding gaseous fuel exiting the combustor jacket to power the fuel pump and/or an oxidizer pump, a flow control device for directing expanded gaseous fuel exiting the turbine through a first outlet and/or a second outlet, a discharge device for exhausting the expanded gaseous fuel exiting the first outlet, a heat exchanger for cooling the expanded gaseous fuel exiting the second outlet with some of the pressurized liquid fuel exiting the fuel pump, and a mixer for combining cooled fuel exiting the heat exchanger with liquid fuel to form mixed liquid fuel to be directed into the fuel pump. Another liquid-fueled rocket engine assembly and related methods are also described.

Abstract: Closed-cycle rocket engine assemblies including a combustor assembly, a combustor jacket, a turbine, a first pump and a mixing chamber are disclosed. The combustor jacket facilitates the transfer of heat from the combustor assembly into a fluid and the turbine is driven by a heated fluid from the combustor jacket. The mixing chamber may include a first inlet to receive a fluid from the turbine, a second inlet to receive a fluid from a first reactant reservoir, and an outlet to deliver a fluid to the first pump. Additionally, the first pump may be coupled to and powered by the turbine and the first pump may be configured to deliver at least a portion of the fluid from the mixing chamber into the combustion chamber of the combustor assembly. Related methods of operating such rocket engine assemblies are also disclosed.

Abstract: A cooling system for a high-speed vehicle may comprise a combustor wall at least partially enclosing a combustor and which is cooled using a coolant circulating in a Brayton cycle. The heated coolant may be expanded in a turbine, transfer heat to a fuel within a heat exchanger, and be compressed by a compressor before returning to the combustor wall. The combustor wall may be capable of withstanding high temperatures, higher than the temperature at which fuel coking may take place. Heat transfer takes place between the coolant and the combustor wall, and between the coolant and the fuel. A method of cooling an engine for a high-speed vehicle is also disclosed.

Abstract: Closed-cycle rocket engine assemblies including a combustor assembly, a combustor jacket, a turbine, a first pump and a mixing chamber are disclosed. The combustor jacket facilitates the transfer of heat from the combustor assembly into a fluid and the turbine is driven by a heated fluid from the combustor jacket. The mixing chamber may include a first inlet to receive a fluid from the turbine, a second inlet to receive a fluid from a first reactant reservoir, and an outlet to deliver a fluid to the first pump. Additionally, the first pump may be coupled to and powered by the turbine and the first pump may be configured to deliver at least a portion of the fluid from the mixing chamber into the combustion chamber of the combustor assembly. Related methods of operating such rocket engine assemblies are also disclosed.

Abstract: The invention is a coolant system for a rocket engine. The rocket engine includes an injector, a fuel supply, an oxidizer supply, a pump for feeding fuel from the fuel supply to the injector, a pump for feeding oxidizer from the oxidizer supply to the injector, a combustor, and a nozzle, the combustor and nozzle forming a combustor and nozzle assembly. The coolant system includes a cooling jacket surrounding the combustor and nozzle assembly. A coolant fluid is supplied to the combustor jacket by a coolant pump that circulates the coolant fluid through the jacket. The coolant fluid vaporizes to cool the combustor and nozzle assembly. A turbine is operatively connected to the coolant pump and is driven by the vaporized coolant from the jacket. A heat exchanger transfers heat from the vaporized fluid to fuel or oxidizer from the supply thereof prior to the feeding of the fuel or oxidizer into the injector.

Abstract: A multi-mode multi-propellant rocket engine capable of operating in a plurality of selected modes. G f ⁡ ( K X + 1 ) ⁢ C * A * ⁢ P 0 = 1 Propellant components may include liquid hydrogen, liquid hydrocarbon, liquid oxygen, liquid fluorine, and liquid air. The liquid oxygen and the liquid air are stored in separate tanks are mixed in a dedicated mixer prior to their injection into the combustion chamber.