A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

A rocket engine pintle injector with optimized spray pattern and with integrated igniter design for providing construction simplicity, throttleable thrust, stop/start/restart capability, optimized operational combustion, and improved ignition combustion stability. A user can start, throttle, and stop the engine by moving the internal concentric injector sleeve forward and backward to cause the fuel/oxidizer to spray out of the pintle head at different flow rates. The concentric igniter can be deployed so that the hot gasses or spark produced are radially projected into the spray of fuel/oxidizer surrounding the igniter. Once the fuel/oxidizer spray has been ignited, the igniter is stopped and retracted to protect the device from the heat of the combustion chamber and is ready for redeployment and restart of the engine as needed. Thus, a versatile, fully integrated, and scalable device can be used to start, throttle, stop, and restart any size rocket engine during any mission phase from launch to return from space.

A flight test system uses a flight termination destruct charge that is configured to overpressurize a pressure vessel in a rocket motor to terminate thrust. The flight termination destruct charge is an electroexplosive detonator arranged on a final burn surface of a propellant contained in the pressure chamber. In a multi-pulse rocket motor, one of the pulses is ignited by the activation of the detonator. The activated detonator is configured to ignite the propellant grain without venting of the gas resulting from the burning of the propellant. Due to the burning of the propellant, the surface area in the pressure vessel is increased which causes increased pressure in the pressure vessel until a critical pressure is reached. When the critical pressure is reached, the rocket motor casing structural capabilities are exceeded. The overpressurized rocket motor casing then ruptures and thrust of the rocket motor is terminated.

An actuation system includes a power drive unit that drives a plurality of drive shafts that each has a first end individually connected to the power drive unit. The power drive unit includes a geartrain including a plurality of individual gears that are offset relative to each other along the gearbox, and a motor-driven rotation of one of the plurality of individual gears drives rotation of the other individual gears, and each of the plurality of individual gears rotates to drive rotation of a respective one of the plurality of drive shafts. A plurality of torque brakes is mounted to the gearbox, the torque brakes being mechanically coupled to respective individual gears and drive shafts. When the torque being passed through is above a predetermined threshold value, a locking of one of the torque brakes stops rotation of all gears simultaneously to maintain positional symmetry in the actuation system.

A turbine pump assembly has a turbine, a centrifugal pump, a passive electrical speed control system, and a pneumatic circuit breaker. The pneumatic circuit breaker has a plurality of elements that are configured to move to a position blocking an outlet duct of the turbine when a flow velocity exceeds a predetermined threshold. A rocket thrust vector control system is also disclosed.

A turbine pump assembly has a turbine, a centrifugal pump, a passive electrical speed control system, and a pneumatic circuit breaker. The pneumatic circuit breaker has a plurality of elements that are configured to move to a position blocking an outlet duct of the turbine when a flow velocity exceeds a predetermined threshold. A rocket thrust vector control system is also disclosed.

A thermally initiated variable venting system may comprise a first linear shape charge (LSC) coupled to a first sensor and a second LSC coupled to a second sensor. An upper apex of the second LSC may be disposed within a lower apex of the first LSC. The output of the system may vary depending on whether the event is fast cook-off (FCO) or slow cook-off (SCO).

A thermally initiated variable venting system may comprise a first linear shape charge (LSC) coupled to a first sensor and a second LSC coupled to a second sensor. An upper apex of the second LSC may be disposed within a lower apex of the first LSC. The output of the system may vary depending on whether the event is fast cook-off (FCO) or slow cook-off (SCO).

A rocket engine pintle injector with optimized spray pattern and with integrated ignitor design for providing construction simplicity, throttleable thrust, stop/start/restart capability, optimized operational combustion, and improved ignition combustion stability. A user can start, throttle, and stop the engine by moving the internal concentric injector sleeve forward and backward to cause the fuel/oxidizer to spray out of the pintle head at different flow rates. The concentric ignitor can be deployed so that the hot gasses or spark produced are radially projected into the spray of fuel/oxidizer surrounding the ignitor. Once the fuel/oxidizer spray has been ignited, the ignitor is stopped and retracted to protect the device from the heat of the combustion chamber and is ready for redeployment and restart of the engine as needed. Thus, a versatile, fully integrated, and scalable device can be used to start, throttle, stop, and restart any size rocket engine during any mission phase from launch to return from space.