Abstract: An apparatus for guiding a transition between flight modes of an electric aircraft is illustrated. The apparatus comprises at least a sensor configured to detect a movement datum of the electric aircraft and a flight controller communicatively connected to the at least sensor, wherein the flight controller is configured to receive the movement datum from the at least a sensor, determine a current flight mode of the electric aircraft as a function of a pilot input and the movement datum, generate a guidance datum as a function of a change in flight mode and the movement datum, and communicate the guidance datum to a pilot indicator in communication with the at least a sensor and flight controller communicatively connected to the at least a sensor.

Abstract: A thrust vector control actuator is provided including a ram, a first plate housed within the ram, a second plate housed within the ram, and a dividing wall housed within the ram. The dividing wall being located between the first plate and the second plate. The dividing wall defines a first chamber within the ram comprising the first plate and a second chamber within the ram comprising the second plate. The actuator also includes an output rod housed within the ram. The output rod having a first end and a second end. The second end is configured to operably connect to an output link. The actuator also includes a load relieving mechanism located within the ram. The load relieving mechanism configured to operatively connect the ram and the output rod. The load relieving mechanism is configured to absorb at least one transient load on the output rod.

Abstract: A solar generator tug is disclosed. The tug can dock with spacecraft to provide power for the spacecraft. Further, the tug may dock with other specialty tugs to form a custom transport system.

Abstract: An electromechanical actuation system (10) for a space vehicle includes a propellant source (18) and at least one turbine (14) operably connected to the propellant source (18) and rotatable by a flow of propellant (16) therefrom. At least one electrical generator (20) is operably connected to the at least one turbine (14) and is configured to convert rotation of the turbine (14) into electrical energy. At least one electromechanical actuator (12) is operably connected to the at least one electrical generator (20) such that electrical energy from the at least one electrical generator (20) drives operation of the at least one electromechanical actuator (12). A method of operating a turbine powered electromechanical actuator (12) for a space vehicle includes rotating the at least one turbine (14) via a flow of propellant (16) therethrough.

Abstract: An electromechanical actuation system (10) for a space vehicle includes a propellant source (18) and at least one turbine (14) operably connected to the propellant source (18) and rotatable by a flow of propellant (16) therefrom. At least one electrical generator (20) is operably connected to the at least one turbine (14) and is configured to convert rotation of the turbine (14) into electrical energy. At least one electromechanical actuator (12) is operably connected to the at least one electrical generator (20) such that electrical energy from the at least one electrical generator (20) drives operation of the at least one electromechanical actuator (12). A method of operating a turbine powered electromechanical actuator (12) for a space vehicle includes rotating the at least one turbine (14) via a flow of propellant (16) therethrough.

Abstract: The invention relates to a device for damping the lateral forces due to jet separation that act on a rocket engine nozzle during a stage of starting or stopping the engine, the engine including at least two identical drive assemblies mounted on the nozzle to take up the lateral forces acting thereon, each drive assembly comprising a first member forming a strut, a second member forming an anchor structure, and an actuator. Each strut comprises means enabling it to act as a rigid strut so long as the lateral forces acting on the corresponding actuator remain below a determined force threshold, and as an element for peak limiting force and for dissipating residual kinetic energy once the forces acting on the actuator exceed the determined force threshold.

Abstract: The present invention relates to a propulsion system of a vertical takeoff and landing aircraft or vehicle moving in any fluid or vacuum and more particularly to a vector control system of the vehicle propulsion thrust allowing an independent displacement with six degrees of freedom, three degrees of translation in relation to its centre of mass and three degrees of rotation in relation to its centre of mass. The aircraft displacement ability using the propulsion system of the present invention depends on two main thrusters or propellers and which can be tilted around pitch is (I) by means of tilting mechanisms and, used to perform a forward or backward movement, can be tilted around roll axis (X) by means of tilting mechanisms and, used to perform lateral movements to the right or to the left and to perform upward or downward movements (Z), the main thrusters being further used to perform rotations around the vehicle yaw axis (Z) and around the roll is (X).

Abstract: An object in motion has a force applied thereto at a point of application. By moving the point of application such that the distance between the object's center-of-mass and the point of application is changed, the object's orientation can be changed/adjusted.

Abstract: The present invention relates to a propulsion system of a vertical takeoff and landing aircraft or vehicle moving in any fluid or vacuum and more particularly to a vector control system of the vehicle propulsion thrust allowing an independent displacement with six degrees of freedom, three degrees of translation in relation to its centre of mass and three degrees of rotation in relation to its centre of mass. The aircraft displacement ability using the propulsion system of the present invention depends on two main thrusters or propellers and which can be tilted around pitch is (I) by means of tilting mechanisms and, used to perform a forward or backward movement, can be tilted around roll axis (X) by means of tilting mechanisms and, used to perform lateral movements to the right or to the left and to perform upward or downward movements (Z), the main thrusters being further used to perform rotations around the vehicle yaw axis (Z) and around the roll is (X).

Abstract: Disclosed is a space propulsion device capable of continuous operation using a solid propellant having high density and excellent handleability.

Abstract: The invention relates to a device for damping the lateral forces due to jet separation that act on a rocket engine nozzle during a stage of starting or stopping the engine, the engine including at least two identical drive assemblies mounted on the nozzle to take up the lateral forces acting thereon, each drive assembly comprising a first member forming a strut, a second member forming an anchor structure, and an actuator. Each strut comprises means enabling it to act as a rigid strut so long as the lateral forces acting on the corresponding actuator remain below a determined force threshold, and as an element for peak limiting force and for dissipating residual kinetic energy once the forces acting on the actuator exceed the determined force threshold.

Abstract: A system and method for supplying thrust to a structure, such as a satellite or spacecraft, for the purposes of station keeping and attitude control of the structure in low-gravity (orbital) and zero-gravity environments. The system includes devices for emitting energy beams and targets impacted by the energy beams to cause ablation of the targets. The beam-emitting devices and targets are adapted to cooperate and cause the structure to selectively undergo translational and/or rotational motion in reaction to the motion of material ablated from the targets. The position, alignment, and/or attitude of the structure can thereby be controlled in a zero or low-gravity environment by selectively emitting the energy beams at the targets.

Abstract: An apparatus and method including a deployable spacecraft mount for electric propulsion is disclosed. A typical apparatus includes a spacecraft body, a deployable element having at least two basic positions including a compact stowed position and a deployed position of the element and an electric thruster disposed on the deployable element where the electric thruster is disposed to mitigate negative plume effects in the deployed position that would be present in stowed position. The deployable element can be a radiator and can optionally be disposed on a second deployable payload module. Further, deployable elements can be selective deployed such that some electric thrusters can be used to assist in transfer orbit while undeployed elements help retain heat.