Abstract: A marine propulsion system includes a controller configured or programmed to perform a rudder angle change control to change a rudder angle of an auxiliary propulsion device by a predetermined angle to one side in a right-left direction of a hull with respect to a forward-rearward direction of the hull so as to move the hull along the forward-rearward direction without rotating the hull when motorized forward-rearward movement is performed to move the hull along the forward-rearward direction by driving the auxiliary propulsion device that is provided to one side of the hull in the right-left direction without generating a thrust from a main propulsion device.

Abstract: A thrust generator is provided for producing thrust to move a spacecraft. The thrust generator includes a housing having a first end and an opposing second end. The first end is associated with a mount for coupling to the spacecraft. The housing further defines a central axis extending through the first end and the second end. The second end defines an annular propulsion outlet. At least one nozzle is positioned proximate the second end. The thrust generator is selectively operable in a first mode in which the thrust generator uses propellant to electrostatically generate thrust via the annular propulsion outlet, and a second mode in which the thrust generator uses propellant to gas-dynamically generate thrust via the at least one nozzle.

Abstract: A flight-capable drone that can optionally comprise: a body; a plurality of propellers each coupled to and having a fixed rotational axis relative to the body, and a plurality of propeller guards. The plurality of propellers operably configured to enable flight along a flight path. During the flight path the plurality of propellers are pitched forward at an angle relative to the flight path. The plurality of propeller guards are fixedly coupled to the body and have a least a part-annular extent. Each one of the plurality of propeller guards positioned radially adjacent a tip of a corresponding one of the plurality of propellers to protect the corresponding propeller from contact. The plurality of propeller guards are configured as an airfoil along at least a portion of the part-annular extent thereof to generate lift during flight along the flight path.

Abstract: A rocket engine is provided with a propellant valve controlling the output of a propellant tank. The valve is configured to induce a vortex in the propellant as it passes the propellant valve. Plural rocket engines may be supplied by a single valve, the valve inducing a vortex in the propellant which is then supplied to plural propellant lines configured to enhance the vortex induced by the valve.

Abstract: The disclosure relates to an electric drive system for an aircraft and to a corresponding operating method. The electric drive system includes a multiplicity of electric thrust generators, wherein each electric thrust generator has an electric motor, and a subsystem group including a multiplicity of subsystems. The thrust generators are apportioned unambiguously between the subsystems so that each subsystem includes two or more of the thrust generators. Furthermore, a control system for operating the drive system is provided, wherein the control system is configured to operate the drive system in such a way that, at least in the event that one of the subsystems of the subsystem group is faulty, an overall yawing moment which is composed of the sum of the yawing moments of the thrust generators of each non-faulty subsystem of the subsystem group essentially disappears.

Abstract: Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

Abstract: A rotorcraft having an aerodynamic device arranged below a rotor, which rotor participates at least in providing lift for the rotorcraft in the air, the rotor being mounted to rotate about a first axis of rotation, the aerodynamic device having a fairing provided with at least one air inlet for enabling a stream of cool air to flow from a region that is situated outside the rotorcraft to another region that is situated inside the rotorcraft; at least at a mouth of the at least one air inlet in the fairing, the aerodynamic device has at least one moving flap that is mounted to move in rotation, the at least one moving flap having at least one degree of freedom of movement in rotation about a second axis of rotation relative to the fairing, and the at least one moving flap orienting itself automatically and passively.

Abstract: The invention is directed to an alignment assembly for changing the relative position of a plate of a pedestal assembly with respect to a processing chamber of a reactor. The alignment assembly is connected at a first end to a riser shaft of the heating assembly and at a second end to a drive shaft. One or more portions of the alignment assembly may be selectively axially rotated or laterally moved change the relative position of the plate with respect to the processing chamber as desired.

Abstract: A hybrid rocket engine is described that achieves stable, highly efficient hybrid combustion by having a core flow of fuel-rich gas generator gases, with the flow being surrounded with an annular injection of oxidizer. The fuel-rich gas serves to vaporize and decompose the oxidizer, such as nitrous oxide, and prepare it for effective, stable combustion. In one embodiment, this is done at the head-end of a combustion chamber. The combustion products can then be expanded through a nozzle to create thrust. The engine can be an upper stage engine that can include modular thrust chambers and an integrated aerospike nozzle. The thrust chambers can be arranged in an array that rings the top of the aerospike nozzle.

Abstract: An assembly is provided for an aircraft propulsion system with an axial centerline. The assembly includes a translating sleeve, a sleeve actuator and a locking apparatus. The sleeve actuator is connected to the translating sleeve. The sleeve actuator is configured to move the translating sleeve along the axial centerline between a stowed position and a deployed position. The locking apparatus includes a lock and a lock actuator. The lock comprises a locking element. The lock actuator is configured to move the locking element between a locked position and an unlocked position. The locking element is configured to lock the translating sleeve in the stowed position when the locking element is in the locked position. A first portion of the locking apparatus axially overlaps the sleeve actuator along the axial centerline. A second portion of the locking apparatus does not axially overlap the sleeve actuator along the axial centerline.

Abstract: A lock system for a component of a thrust reverser actuation system (“TRAS”), comprising a lock member translatable between a locked position, in which the lock member prevents movement of the thrust reverser component to deploy the thrust reverser actuation system, and an unlocked position, in which the lock member allows movement of the thrust reverser component to deploy the thrust reverser actuation system. The actuator may further comprise a screw shaft and a nut translatable along the screw shaft and operatively connected to the lock member. Rotational movement of one of the screw shaft and the nut causes the nut to translate along the screw shaft and, in turn, the lock member to move between its locked position and its unlocked position.

Abstract: The invention relates to a combustion gas discharge nozzle for a rocket engine including a stationary part and a moving part extending from the stationary part, the moving part made using flaps positioned downstream from the stationary part and forming an extension of the nozzle, the nozzle including a sealing device extending between the fixed part and the moving part in the form of a flexible membrane withstanding a local temperature of the combustion gases at the nozzle outlet and connecting the end of the stationary part to a border of the flaps or petals forming the moving part, the flexible membrane forming an annular tubing, the sealing device being provided with a duct for injecting gas at the flexible membrane between the stationary part and the moving part extending the nozzle.

Abstract: A flow vectoring turbofan engine employs a fixed geometry fan sleeve and core cowl forming a nozzle incorporating an asymmetric convergent/divergent (con-di) and/or curvature section which varies angularly from a midplane for reduced pressure in a first operating condition to induce flow turning and axially symmetric equal pressure in a second operating condition for substantially axial flow.

Abstract: A thrust reverser system deployable from a core cowl is provided. The thrust reverser system may comprise one or more doors that are configured to overlap to reduce leakage and increate reverse thrust while deployed. The doors may be configured to vanes that are configured to split and/or direct airflow.

Abstract: A blocker door actuation system for use in an aircraft thrust reverser is provided. The blocker door actuation system may comprise a rack, a gear, a gear housing, a screw shaft, and a link. The blocker door actuation system may be mounted to a translating sleeve in an aircraft nacelle. In response to the thrust reverser being activated and the translating sleeve moving aft, the blocker door actuation system may move the blocker door from a stowed position to a deployed position in a fan air duct.

Abstract: A thrust-vectoring nozzle is disclosed that includes a cylindrical case and a flow director. The flow director includes a cylindrical ring having a dimension to fit within the cylindrical case. The cylindrical ring has an inner wall and an outer wall and a plurality of fixed vanes are coupled to the inner wall of the ring. The flow director is configured to rotate about first and second mutually perpendicular axes.

Abstract: In one embodiment, a nozzle of a gas turbine engine may be provided having a coanda injector and a fluidic injector which operate together to provide for a change in exhaust flow direction. The fluidic injector may be coincident with or downstream of the coanda injector and both may be used in high pressure ratio operations of the nozzle. The fluidic injector may be positioned opposite the coanda injector and, when activated, may provide for a region of separated flow on the same side of the nozzle as the fluidic injector. The coanda injector may provide additional momentum to an exhaust flow flowing through the nozzle and may encourage the flow to stay attached on the coanda injector side of the nozzle.

Abstract: The present invention's nozzle is especially suitable for use in a shipboard air-discharge maneuverability device known as a “bow thruster.” As typically embodied, the inventive nozzle includes a nozzle wall and an arc-shaped vane. The nozzle wall has a circular nozzle inlet and an elliptical nozzle outlet, and is configured so that the geometric major axis of the nozzle outlet is horizontal. The arc-shaped vane: is upwardly bowed in the nozzle wall's transverse direction; joins the nozzle wall along its two horizontally opposite sides in the nozzle wall's longitudinal direction; extends most or all of the longitudinal distance between the nozzle inlet and the nozzle outlet; has a transversely intermediate section that, taken in the transverse direction, is generally uniform in thickness; has two transversely lateral sections that, taken in the transverse direction, gradually thicken toward the two respective joints of the arc-shaped vane with respect to the nozzle wall.

Abstract: A detonation thrust-producing device includes an explosive located in a recess in an external surface of a body. Detonation of the explosive expels material out of the recess, providing thrust to the body in an opposite direction. A mass, such as a metal disk, may be placed blocking or covering the external opening. The body may be a part of a vehicle, such as an airborne projectile. The thrust-producing device may include multiple detonation motors arrayed around the body, capable of being individually or multiply detonated. Such thrust-producing devices may be used for attitude adjustment, steering, or other control of the flight of the projectile or other air vehicle. The detonation thrust-producing devices have the advantage of a faster-response time than propellant-based devices, and do not need the nozzles that are used with many propellant-based devices.

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: A hot-gas thruster is actuated using a relatively small electric motor. The hot-gas thruster includes a pressure assisted pilot shaft to keep electric power demand to only a few hundred watts peak and only tens of watts on average, while exhibiting relatively fast response times.

Abstract: The exhaust system is located on each nacelle of a tiltrotor aircraft. The exhaust system includes a vector nozzle that is selectively rotatable in relation to each nacelle in order to achieve certain performance objectives. The vector nozzle can be oriented to provide maximum flight performance, reduce infrared (IR) signature, or even to reduce/prevent ground heating.

Abstract: A dual redundant actuation control system for controlling a plurality of actuators for positioning a plurality of moveable aircraft components. The actuation control system includes a component controller. The component controller includes two component control channels. Each of the plurality of actuators is electrically connected to each of the two component control channels such that either of the two component control channels may control any or all of the plurality of actuators.

Abstract: A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel.

Abstract: A by-pass turbojet including a thrust reverser is disclosed. The thrust reverse includes a deflector device which deflects all or part of the primary stream gas in such a manner that the deflected primary stream gas encounters the secondary stream, thereby reducing the injection speed of the secondary stream in an aft direction, and thus generating the thrust-reversal effect. The secondary stream escapes from the aft end of the nacelle. In the thrust-reversal position, the deflector device is inscribed radially substantially within the section of the primary nozzle cowl at the aft end of the nacelle. Such a thrust reverser is particularly simple in design, inexpensive, and makes it possible to avoid having moving parts present on the outside portion of the nacelle, thus simplifying the design of the turbojet.

Abstract: A vectoring nozzle with external actuation generates thrust vectoring by applying mechanical or fluidic actuation, or both, on the nozzle deck, external sidewalls, and/or air vehicle aft body to produce changes in the aft body flowfield and/or exhaust plume. An external mechanical sidewall may be integrated into a nozzle deck or side walls without the need for engine bleed to supply fluid injectors. An external fluidic vectoring system uses injectors or plasma devices located aft of the nozzle exit to vector the exhaust plume with no external moving parts. Elements of both mechanical and fluidic systems may be combined for a given application.

Abstract: A nozzle device defines a passageway including an outlet to discharge working fluid to produce thrust. This device includes a vectoring mechanism having three or more vanes pivotally mounted across the passageway and a linkage pivotally coupling the vanes together. This linkage includes a first arm fixed to a first one of the vanes to pivot therewith about a first pivot axis, a second arm and a third arm fixed to a second one of the vanes to pivot therewith about a second pivot axis, and a fourth arm fixed to a third one of the vanes to pivot therewith about a third pivot axis. A first connecting link pivotally couples the first arm and the second arm together, and a second connecting link pivotally couples the third arm and the fourth arm together. The relative angular positioning of the arms with respect to the corresponding pivot axes and/or the arm links can be varied to define different vectoring schedules with the mechanism linkage.

Abstract: A hot-gas thruster is actuated using a relatively small electric motor. The hot-gas thruster includes a pressure assisted pilot shaft to keep electric power demand to only a few hundred watts peak and only tens of watts on average, while exhibiting relatively fast response times.

Abstract: An outlet device for a jet engine comprises a number of fixed ducts, each with a gas intake and a gas outlet for conducting a gas from the jet engine, at least two of the outlets of said gas ducts open in different directions, and a gas distribution arrangement is arranged at the gas intakes for selective distribution of the gas to the ducts.

Abstract: One embodiment of the present invention is a unique thrust augmented gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for thrust augmented gas turbine engines. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Abstract: A method for exhausting gas from an aircraft engine assembly is provided. The method includes coupling a first exhaust duct in fluid communication only to a first engine. The first exhaust duct includes a primary outlet and a secondary outlet. The method further includes coupling a second exhaust duct in fluid communication only to a second engine. The second exhaust duct includes a primary outlet and a secondary outlet. The method also includes aligning a portion of the first engine secondary outlet concentrically within a portion of the second engine secondary outlet.

Abstract: Improved rocket nozzle designs for vehicles with nozzles embedded in or protruding from surfaces remote from the desired thrust axis. The nozzle configurations are for rocket vehicles where the nozzles are not located at the optimal thrust axis of the vehicle. Two examples include nozzles located on the forward end of the vehicle (also called tractor nozzles) and attitude control nozzles located on the periphery of the vehicle. More particularly, the disclosed nozzle shapes enhance the axial thrusts and/or maneuver torques on the vehicle. These unconventional nozzle shapes improve vehicle performance.

Abstract: The invention relates to a system for controlling a plurality of actuators (15) that can displace a mobile panel (13, 17) pertaining to a nacelle (1) of an aircraft, said system comprising at least two motors (16) that can drive the actuators (15). The system also comprises two separate control units (33, 35), each unit being configured in such a way as to control and feed at least one motor that is not fed or controlled by the other control unit. The invention also relates to a nacelle comprising such a system.

Abstract: A swept fan ramp for a pivot door thrust reverser includes a cylindrical unit with an elliptical flared portion, a side portion, and a rectangular flared portion. The elliptical flared portion is operable to reduce drag on airflow in a reverse direction and is coupled to the top center and bottom center of the circumference of the cylindrical unit. The side portion is operable to reduce side spillage airflow and is coupled to the central left side and central right side of the aft circumference of the cylindrical unit. The rectangular flared portion is operable to promote separation of airflow into an upper airflow path and a lower airflow path, is coupled with the aft circumference of the cylindrical unit, and is connected to the elliptical flared portion and the side portion.

Abstract: A non-invasive system, method, and apparatus for control input prediction and state verification of an aircraft's fluidic vectoring exhaust is disclosed. The control system derives a desired vector state, then predicts and sets the fluidic injection input required to produce the desired vector state. A vectored state verification routine is used to determine the resulting vector state for feedback to the control system.

Abstract: An exhaust vane disposed in a divergent section of an aircraft gas turbine engine exhaust nozzle is sideways pivotable about a vane pivot axis. The nozzle vane pivot axis may be centrally located at an unvectored nozzle throat. Transversely spaced apart upper and lower tips of the exhaust vane may be incorporated to sealingly engage a nozzle outer wall along upper and lower surfaces of the outer wall of the nozzle. The exhaust vane has flat or contoured vane sidewalls and contoured vane sidewalls may be concave. The exhaust vane may have transversely biased apart upper and lower vane sections extending transversely inwardly from the upper and lower tips of the exhaust vane respectively. The exhaust vane may be articulated having upstream and downstream sections separately sideways pivotable about the vane pivot axis and a second pivot axis downstream of and parallel to the vane pivot axis respectively.

Abstract: The present invention relates to an air compression type engine for aviation. The conventional aircraft engine is complicated in structure, consumes expensive aero fuel, and produces extremely loud noise during operation. Therefore, the present invention employs a turbo-charged air compressor to generate high temperature and high pressure gas in the pressure chamber, and uses the reaction thrust force generated by ejecting the compressed air through the nozzle to make the aircraft vertically take off/land, suspend in the air and fly forwardly. Benefited from the simple structure, the manufacture cost of the aircraft engine can be dramatically decreased. Vertically raising and landing the aircraft can be achieved by changing the ejection directed of the compressed air in the pressure chamber. The commonly used gasoline or diesel, which is cost saving and capable of combusting with a high combustion value, can be used instead of the expensive aviation kerosene. Further, the generated noise is quite small.

Abstract: An exhaust nozzle (10) for a gas turbine engine is described for vectoring a flow of exhaust gases issuing therefrom at an angle to a centre line (8) of the engine. The exhaust nozzle (10) comprises a plurality of radially outer flaps (25) and radially inner flaps (26) which are circumferentially disposed around the engine. The radially outer flaps (25) and radially inner flaps (26) alternately overlap each other to define a path through which the exhaust gases flow. The radially outer flaps (25) can be moved asymmetrically about the engine centre line (8). The radially inner flaps (26) are maintained in sealing contact with the outer flaps (25) by the flow of exhaust gases passing therethrough. The radially inner flaps (26) are divided into a number of triangular sections (29) which renders the flaps (26) torsionally flexible so that they can twist along their lengths to maintain sealing contact with adjacent radially outer flaps (25) when the outer flaps (25) are moved asymmetrically.

Abstract: A propulsion system includes a canted multinozzle plate, which has a multitude of small nozzles angled (not perpendicular) to major surfaces of the multinozzle grid plate. The multinozzle plate may be a cylindrical section or plate, and the multitude of nozzles may be substantially axisymmetric about the cylindrical plate. The propulsion system includes a pressurized gas source which may be placed either forward or aft of the multinozzle grid plate. The propulsion system may have a conical insert, an internal flow separator cone, to aid in changing directions of flow from the pressurized gas source, to divert the flow through the multiple nozzles.

Abstract: A nozzle device defines a passageway including an outlet to discharge working fluid to produce thrust. This device includes a vectoring mechanism having three or more vanes pivotally mounted across the passageway and a linkage pivotally coupling the vanes together. This linkage includes a first arm fixed to a first one of the vanes to pivot therewith about a first pivot axis, a second arm and a third arm fixed to a second one of the vanes to pivot therewith about a second pivot axis, and a fourth arm fixed to a third one of the vanes to pivot therewith about a third pivot axis. A first connecting link pivotally couples the first arm and the second arm together, and a second connecting link pivotally couples the third arm and the fourth arm together. The relative angular positioning of the arms with respect to the corresponding pivot axes and/or the arm links can be varied to define different vectoring schedules with the mechanism linkage.

Abstract: Method for the computer-assisted determination of an optimum-fuel control of nozzles according to a control instruction b=Ax. A defined matrix transformation of starting constraints for the mass flow of the nozzles and of the minimization criterion thereby takes place in a computer-assisted manner, a data processing representation of a geometric description of the matrix-transformed starting constraints, a computer-assisted determination of limiting point sets of the geometric description of the starting constraints through a computer-assisted geometric search procedure in the vector space and the application of the matrix-transformed minimization criterion to the points of the limiting point sets.

Abstract: A first trailing edge portion of a scarfed jet engine exhaust nozzle aft of a second trailing edge portion relative to a central axis of an associated exhaust duct causes an automatic nozzle-pressure-ratio responsive transverse deflection of the associated exhaust flow away from the first trailing edge portion. When offset from both the center of gravity (CG) and the central longitudinal axis of an aircraft, at a relatively low nozzle pressure ratio, e.g. during takeoff, the thrust vector from the exhaust flow acts relatively close to the CG, whereas at a relatively high nozzle pressure ratio, e.g. during relatively high-speed cruise, the scarfed exhaust nozzle deflects the exhaust flow so that the resulting thrust vector is relatively parallel to the path of the aircraft. With the final portion of the exhaust duct skewed, the primary axis of the jet engine can be relatively parallel to the path of the aircraft.

Abstract: The present invention provides an actuating apparatus for moving relatively rotatable components. In one embodiment the apparatus comprises: a bearing (30) including first (34) and second (36) relatively rotatable bearing elements for mounting respective first (18) and second (12) rotatable components for relative rotation about a common axis; a pneumatic or hydraulic actuator having a pair of co-acting parts (44, 54), including a first part (54) fixed with respect to the said first rotatable element and a second part (44) fixed with respect to the said second rotatable element; and, means (110) for energising and/or de-energising said actuator means to effect movement of the said co-acting parts and thereby relative rotation of the said first and second bearing elements about the said bearing axis.

Abstract: A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel.

Abstract: A vectoring nozzle with external actuation generates thrust vectoring by applying mechanical or fluidic actuation, or both, on the nozzle deck, external sidewalls, and/or air vehicle aft body to produce changes in the aft body flowfield and/or exhaust plume. An external mechanical sidewall may be integrated into a nozzle deck or side walls without the need for engine bleed to supply fluid injectors. An external fluidic vectoring system uses injectors or plasma devices located aft of the nozzle exit to vector the exhaust plume with no external moving parts. Elements of both mechanical and fluidic systems may be combined for a given application.

Abstract: According to the invention, under braking of an aircraft (AC) with reverse thrust, that (those) engine(s) (Mc) of said aircraft (AC) the associated control lever (4c) of which is (are) not in a reverse thrust position (I) is (are) set to low idle speed.

Abstract: An engine assembly includes a gas-turbine engine having a tailcone portion and a bypass duct, a rocket engine combustion assembly located at the tailcone portion of the gas-turbine engine, and a movable nozzle segment subassembly that is selectively engageable with the gas-turbine engine bypass duct in an open position and with the rocket engine combustion assembly in a closed position.

Abstract: Method for the computer-assisted determination of an optimum-fuel control of nozzles according to a control instruction b=Ax. A defined matrix transformation of starting constraints for the mass flow of the nozzles and of the minimization criterion thereby takes place in a computer-assisted manner, a data processing representation of a geometric description of the matrix-transformed starting constraints, a computer-assisted determination of limiting point sets of the geometric description of the starting constraints through a computer-assisted geometric search procedure in the vector space and the application of the matrix-transformed minimization criterion to the points of the limiting point sets.

Abstract: A ducted air power plant, comprising a motor driven fan (7) situated in a duct (4), the fan (7) having an air intake side and in operation providing a high pressure air stream in the duct, and the fan being located adjacent air splitter mechanism (18), the air splitter mechanism (18) being arranged to divert the air stream into two or more subsidiary streams for delivery to respective jet nozzles (9) of the plant. The plant may be used in a vehicle such as an aircraft in order to provide a vertical take-off and hover capability as well a level flight power source.

Abstract: An ejection seat having pitch, roll, and yaw control provided by three separate rocket motors where each rocket motor has a fixed nozzle and the entire rocket motor is rotated about a single axis corresponding to the minimum principal moment of inertia of the rocket. Actuation for each rocket motor is by means of a hydraulic rack and pinion actuator. Power for the hydraulic actuators is provided by a unique hydro-pneumatic amplifier that converts stored gas energy into pressurized hydraulic fluid. The high pressure hydraulic fluid is directed through conventional servo valves into the appropriate actuators to provide main, roll, pitch, and yaw thrust as required to achieve upright orientation and vertical flight.