Various embodiments of a vortex thruster system is described herein that is configured to create at least three discrete thrust levels. In some embodiments, the vortex thruster system is configured to decompose a monopropellant and deliver the decomposed monopropellant into a vortex combustion chamber for generating various thrust levels. In some embodiments, the vortex thruster system includes a secondary propellant valve configured to deliver a secondary propellant into the vortex combustion chamber containing decomposed monopropellant to create a high thrust level. Related systems, methods, and articles of manufacture are also described.

A readily combustible portion (110) includes a readily combustible exposed surface (111) that is exposed to a flow channel (CA). A combustion-resistant portion (140), which comprises a material that is more resistant to combustion than the readily combustible portion (110), covers an outer surface of the readily combustible portion (110) on the opposite side from the readily combustible exposed surface (111) in a direction orthogonal to a length direction parallel to a direction in which a hybrid rocket is propelled. The combustion-resistant portion (140) includes a thick portion (120) that serves as a stopper that prevents peeling of the readily combustible portion (110) from the combustion-resistant portion (140) in a direction from a starting end surface (100a) toward a terminating end surface (100b).

A readily combustible portion (110) includes a readily combustible exposed surface (111) that is exposed to a flow channel (CA). A combustion-resistant portion (140), which comprises a material that is more resistant to combustion than the readily combustible portion (110), covers an outer surface of the readily combustible portion (110) on the opposite side from the readily combustible exposed surface (111) in a direction orthogonal to a length direction parallel to a direction in which a hybrid rocket is propelled. The combustion-resistant portion (140) includes a thick portion (120) that serves as a stopper that prevents peeling of the readily combustible portion (110) from the combustion-resistant portion (140) in a direction from a starting end surface (100a) toward a terminating end surface (100b).

Various embodiments of a vortex hybrid motor are described herein. In some embodiments, the vortex hybrid motor may include a combustion zone defined by a fuel core and/or motor housing. The combustion zone may include an upper zone and a central zone that each contribute to thrust created by the vortex hybrid motor. In some embodiments, an injection port configuration is described that includes a proximal injection port that may be controlled for modulating a delivery of an amount of oxidizer for adjusting an oxidizer-to-fuel ratio. In some embodiments, a fuel core configuration is described that provides radially varying gradients of fuel in order to achieve desired thrust profiles. In some embodiments, the fuel core may include a support structure and/or a proximal end of a nozzle of the vortex hybrid motor may extend into the fuel core.

Various embodiments of a vortex hybrid motor are described herein. In some embodiments, the vortex hybrid motor may include a combustion zone defined by a fuel core and/or motor housing. The combustion zone may include an upper zone and a central zone that each contribute to thrust created by the vortex hybrid motor. In some embodiments, an injection port configuration is described that includes a proximal injection port that may be controlled for modulating a delivery of an amount of oxidizer for adjusting an oxidizer-to-fuel ratio. In some embodiments, a fuel core configuration is described that provides radially varying gradients of fuel in order to achieve desired thrust profiles. In some embodiments, the fuel core may include a support structure and/or a proximal end of a nozzle of the vortex hybrid motor may extend into the fuel core.

Various embodiments of a vortex thruster system is described herein that is configured to create at least three discrete thrust levels. In some embodiments, the vortex thruster system is configured to decompose a monopropellant and deliver the decomposed monopropellant into a vortex combustion chamber for generating various thrust levels. In some embodiments, the vortex thruster system includes a secondary propellant valve configured to deliver a secondary propellant into the vortex combustion chamber containing decomposed monopropellant to create a high thrust level. Related systems, methods, and articles of manufacture are also described.

Various embodiments of a vortex thruster system is described herein that is configured to create at least three discrete thrust levels. In some embodiments, the vortex thruster system is configured to decompose a monopropellant and deliver the decomposed monopropellant into a vortex combustion chamber for generating various thrust levels. In some embodiments, the vortex thruster system includes a secondary propellant valve configured to deliver a secondary propellant into the vortex combustion chamber containing decomposed monopropellant to create a high thrust level. Related systems, methods, and articles of manufacture are also described.

A solid rocket motor is described that includes a solid propellant section, a nozzle, and a source of monopropellant, such as liquid monopropellant. The monopropellant is used to control various operational parameters of the solid rocket motor, such as thrust vector control, roll control, extinguishment of the motor, and cooling of the nozzle and/or nozzle throat. The nozzle and the nozzle throat can be an integrated, single piece assembly that facilitates re-use of the nozzle.

A solid rocket motor is described that includes a solid propellant section, a nozzle, and a source of monopropellant, such as liquid monopropellant. The monopropellant is used to control various operational parameters of the solid rocket motor, such as thrust vector control, roll control, extinguishment of the motor, and cooling of the nozzle and/or nozzle throat. The nozzle and the nozzle throat can be an integrated, single piece assembly that facilitates re-use of the nozzle.

An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.