A nozzle assembly according to an exemplary aspect of the present disclosure includes, among other things, a nozzle including a throat section. The nozzle further includes a ball portion of a ball and socket joint. The assembly further includes a vehicle including a socket portion of the ball and socket joint. The nozzle is mounted to the vehicle and the ball portion is received at least partially in the socket portion. A flow regulator is arranged adjacent the throat section and configured to regulate a flow of fluid through the throat section. The flow regulator is attached to the nozzle upstream of the throat section. An actuator is attached to the nozzle, and the actuator is configured to selectively rotate the nozzle via the ball and socket joint about a first axis normal to a longitudinal axis of the vehicle. A rocket and method are also disclosed.

A combustion chamber structure for a rocket engine includes a hot gas wall (12) that surrounds a combustion chamber (40) and has a plurality of first coolant channels (50) and a plurality of second coolant channels (52). The plurality of first (50) and second (52) coolant channels extend from a first longitudinal end (16) of the hot gas wall (12) to a second longitudinal end (18) of the hot gas wall (12) opposite to the first longitudinal end (16). The combustion chamber structure (10) further comprises a first manifold (20) forming a first coolant chamber (30) and a second manifold (22) forming a second coolant chamber (32) being fluidly separated from the first coolant chamber (30). The first (20) and second (22) manifolds are provided at the first longitudinal end (16) of the hot gas wall (12) and extend in a circumferential direction of the hot gas wall (12). The first coolant chamber (30) is fluidly connected to each of the plurality of first coolant channels (50) and the second coolant chamber (32) is fluidly connected to each of the plurality of second coolant channels (52).

Nozzle of a spacecraft engine, comprising a nozzle body extending along a main direction, from a proximal end secured to the engine of the spacecraft, and a free distal end, said nozzle body having a plurality of stiffeners extending from an outer surface of the nozzle body, radially with respect to the main direction, said nozzle comprising a thermal insulation system, comprising at least one strip of insulating material disposed so as to surround the outer surface of the nozzle body over at least part of the height of said nozzle body-, said thermal insulation system comprising a plurality of holding elements, each holding element being positioned so as to surround said strips of insulating material, and being disposed between two stiffeners of the nozzle body.

Nozzle of a spacecraft engine, comprising a nozzle body extending along a main direction, from a proximal end secured to the engine of the spacecraft, and a free distal end, said nozzle body having a plurality of stiffeners extending from an outer surface of the nozzle body, radially with respect to the main direction, said nozzle comprising a thermal insulation system, comprising at least one strip of insulating material disposed so as to surround the outer surface of the nozzle body over at least part of the height of said nozzle body-, said thermal insulation system comprising a plurality of holding elements, each holding element being positioned so as to surround said strips of insulating material, and being disposed between two stiffeners of the nozzle body.

A rocket engine nozzle manufacturable and applicable to tactical missile designs includes an aerospike having a plurality of airfoil fins distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The airfoil fins and an inner perimeter of the combustion chamber define a plurality of apertures which choke an airflow exiting the combustion chamber and cause the airflow to expand supersonically along the airfoil fins. The aerospike rocket engine nozzle requires less machine precision and achieves packing benefits over conventional bell and aerospike nozzle geometries. The configuration of the aerospike rocket engine nozzle also removes the producibility and heating constraints typically encountered with conventional aerospike nozzles in tactical missile applications while improving thrust performance of the rocket engine across a wide range of altitudes.

A rocket engine nozzle manufacturable and applicable to tactical missile designs includes an aerospike having a plurality of airfoil fins distributed around a central longitudinal axis of a rocket engine combustion chamber. The aerospike is integrated on an exit plane at an exit end of the combustion chamber. The airfoil fins and an inner perimeter of the combustion chamber define a plurality of apertures which choke an airflow exiting the combustion chamber and cause the airflow to expand supersonically along the airfoil fins. The aerospike rocket engine nozzle requires less machine precision and achieves packing benefits over conventional bell and aerospike nozzle geometries. The configuration of the aerospike rocket engine nozzle also removes the producibility and heating constraints typically encountered with conventional aerospike nozzles in tactical missile applications while improving thrust performance of the rocket engine across a wide range of altitudes.

A modular reusable catalyst system is provided for a bi-propellant rocket engine and adaptable for a monopropellant, wherein a hydrocarbon fuel is combined with one or more catalysts, and can provide an additional convergent-divergent flow in combination with a thruster.

A modular reusable catalyst system is provided for a bi-propellant rocket engine and adaptable for a monopropellant, wherein a hydrocarbon fuel is combined with one or more catalysts, and can provide an additional convergent-divergent flow in combination with a thruster.

The combustor for a rocket engine includes a combustion room configured to cause a combustion reaction between a fuel and an oxidant, an injector configured to inject the fuel and the oxidant into the combustion room, and a nozzle skirt configured to inject combustion gas generated by the combustion reaction to an outside, and an inertance increasing portion configured to increase an equivalent inertance in a vibration equivalent circuit of the combustor for the rocket engine.

An annular ablative gas blocking device provides for automatic altitude compensation of a rocket engine exhaust plume. The nozzle is over expanded at low level launch altitudes and near optimally expanded at the highest altitude at the terminal burnout or staging altitude of the rocket engine. The ablative gas blocking device in the nozzle exit mitigates low altitude launch effects of an over expanded nozzle and inhibits external atmospheric air entrance into the nozzle at launch. The gas blocking means ablatively erodes away from plume impingement as the rocket ascends in a pre-programmed manner to achieve optimum area expansion ratio at all altitudes.