Abstract: The invention relates to a rocket engine nozzle comprising a system for controlling jet separation of the flow in the nozzle, wherein said control system exhibits a plurality of separation triggering elements (5, 10) arranged in such a way as to generate, from mutually spaced initiation points (9), distinct zones (6) of jet separation, so as to form a three-dimensional separation of the flow. The flow control system can exhibit at least two triggering elements (5, 10).

Abstract: A method and apparatus for augmenting thrust in a rocket traveling through atmospheric gas. Rocket motor designs are provided where a throat(s) from one or more rocket motors eject high-speed primary exhaust gas in a configuration which peripherally surrounds an outlet for induced, secondary gas. The secondary gas is mixed with the jet of primary exhaust gas to add momentum, and therefore thrust. Either expansion deflection or plug type rocket discharge nozzles can be utilized. In one embodiment, a thrust augmentation of over one hundred percent is achieved. In another embodiment, a plurality of rocket motor assemblies each containing a thrust augmenting rocket motor design is affixed to a rocket body. Such rocket motors enhance rocket thrust performance, and enables more efficient payload to rocket motor selection, or, alternatively, allows higher loads to be carried with the same amount of thrust.

Abstract: The present invention reveals a method and apparatus for controlling the effective area and thrust vector angle of a fluid flow. In one embodiment, the fluid flow is controlled in an advanced, high aspect ratio, complex aperture geometry nozzle using asymmetric injection into the subsonic portion of the fluid flow. The present invention vectors the primary flow by partially blocking the flow with an opposed flow across the flow field. A fluidic flow field is defined in a flow container that directs a pressurized, primary fluidic flow from the container towards an exit of the container. A nozzle may cooperate with the exit of the flow container to control the fluidic flow as it exits the flow container. One or more injectors associated with the container are proximate to the effect throat of the primary flow while other are located downstream of to introduce an opposing fluidic flow that interacts with the primary fluidic flow.

Abstract: According to one embodiment of the invention, a system for altering a fluid flow includes a nozzle having a fluid flow and including a converging portion, a diverging portion downstream of the converging portion, and a throat coupling the converging portion to the diverging portion, at least one port located in a wall of the nozzle and angled with respect to the fluid flow, and at least one pulse detonation device operable to inject a plurality of detonation waves in a pulsed manner through the port and into the fluid flow. The pulsed detonation waves operate to alter the fluid flow.

Abstract: The present invention comprises a rocket motor nozzle that replaces the fixed-wall throat with a “wall” created by injecting gas radially into the nozzle. By injecting gas into the flow of combustion products that are going through the nozzle, this will deflect the combustion product flow, restricting and accelerating such flow, just as a fixed-wall does in a standard nozzle. However, by using the “gas-wall” described herein, no erosion will result at the restriction area.

Abstract: A device for controlling the direction of flow of a primary fluid includes one or more injectors oriented to inject a secondary fluid against the direction of flow of the primary fluid. The injector is formed by drilling or otherwise forming a hole at an angle to the surface of one or more sidewalls of an engine nozzle or other device. A feedback controller regulates the amount and duration of the secondary fluid injection to achieve the commanded attitude or attitude rate. The controller is coupled to one or more plenums attached to the sidewall(s). The plenums can be arranged to deliver secondary fluid to one or more of the injectors. Secondary fluid delivery to each plenum can be controlled independently to control the flow of the primary fluid in one or more directions. The device can be used to provide thrust vectoring in an aircraft or other type of vehicle, as well as other applications where it is desired to control the direction of a primary fluid.

Abstract: According to one embodiment of the invention, a system for altering a fluid flow includes a nozzle having a fluid flow and including a converging portion, a diverging portion downstream of the converging portion, and a throat coupling the converging portion to the diverging portion, at least one port located in a wall of the nozzle and angled with respect to the fluid flow, and at least one pulse detonation device operable to inject a plurality of detonation waves in a pulsed manner through the port and into the fluid flow. The pulsed detonation waves operate to alter the fluid flow.

Abstract: A rocket engine (10) generates a flow of hot propulsion fluid through a nozzle (14N) for generating propulsive thrust along a thrust vector 8. Hybrid exhaust gas generators (36,38) have their exhausts (44,54) through the side of the nozzle. Each gas generator includes a fuel grain (46,56) and a source of oxidizer (16,40,50). The fuel grain is kept hot by either or both (a) direct radiation or conduction from the hot propulsion fluid, or (b) by a trickle of oxidizer. The fuel grain can thus react quickly to a substantial flow of oxidizer.

Abstract: A rocket engine (10) generates a flow of hot propulsion fluid through a nozzle (14N). Hybrid exhaust gas generators (36,38) have their exhausts (44,54) through the side of the nozzle. Each gas generator includes a fuel grain (46,56) and a source of oxidizer (16,40,50). The fuel grain is kept hot by either or both (a) direct radiation or conduction from the hot propulsion fluid, or (b) by a trickle of oxidizer. When the thrust vector is to be modified, the appropriate one of the hybrid gas generators receives a flow of oxidizer, and the resulting exhaust gas is injected through the side of the nozzle.

Abstract: A stream of primary gas flowing through a bi-stable thrust vectoring nozzle becomes attached to a first or second surface extending downstream of the nozzle, each surface incorporating one or more control ports for controlling to which surface the stream is attached, wherein relative to the longitudinal axis of the nozzle, the angle of discharge from the first surface is substantially different from the angle of discharge from the second surface, preferably with the first surface substantially aligned with the longitudinal axis of the nozzle. In one embodiment, a plurality of nozzles are arranged with the respective first surfaces substantially aligned with the longitudinal axis of the nozzle combination and each of the respective second surfaces arranged to laterally deflect a respective portion of the stream of primary gas in a respective direction along each of two orthogonal lateral axes.

Abstract: In accordance with the present invention, an aerodynamic turbo engine nacelle, disposable about an engine having an engine face, for mitigating boundary layer separation of intake airflow comprises an engine inlet. The engine inlet is provided with an interior surface and a curved portion. The curved portion having a leading edge. The geometry of the engine inlet is such that the ratio of the circular area defined by the diameter defined by the leading edge of the curved portion to the minimum circular area defined by the interior surface of the engine inlet is less than 1.33. The engine nacelle is further provided with a pressurized fluid injecting device for injecting pressurized fluid at the engine inlet in a direction generally parallel to intake airflow in response to sensed airflow conditions.

Abstract: An apparatus and method for varying the effective cross sectional area of an opening through a fixed geometry nozzle provides a fluidic cross flow with an injector incorporated in the throat of the nozzle 68 proximate to the subsonic portion of the flow through the nozzle. One or more injectors 76 are directed at an injector angle opposed to the subsonic portion of the flow. The opposed cross flow from injector 76 interacts with a primary flow 14 through the nozzle 68 to partially block the nozzle's opening 72, thereby effectively decreasing the cross sectional area of opening 72 in a jet engine 42. A plurality of cross flows proximate to a nozzle's throat 70 permits effective afterburner 64 operations even with a fixed geometry nozzle by allowing throttling of the primary flow 14. Further, variations in the cross flow's mass flow characteristics or injection angle can allow vectoring of the primary flow 14.

Abstract: An apparatus and method for varying the effective cross sectional area of an opening through a nozzle provides a controller 78 that pulses a fluidic cross flow through an injector associated with the nozzle 68 proximate to the opening of the nozzle. The controller 78 directs the injector 76 to provide a pulsed cross flow having a predetermined frequency, amplitude, or wave form. The pulsed cross flow interacts with a primary flow 14 through the nozzle 68 to partially block the nozzle's opening 72, thereby effectively decreasing the cross sectional area of opening 72 in a jet engine 42. A plurality of pulsed cross flows proximate to a nozzle's throat 70 permits effective afterburner 64 operations even with a fixed geometry nozzle by allowing throttling of the primary flow 14. Further, variations in the cross flow's orientation pulse frequency, amplitude or wave form, or the cross flow's mass flow characteristics can allow vectoring of the primary flow 14.

Abstract: A turning vane arrangement (30) for an IR suppressor having a duct (32) adapted for receiving a primary flow PF of engine exhaust. The turning vane arrangement (30) is operative to direct a pressurized secondary flow PSF of low temperature gaseous fluid into the primary flow of engine exhaust for reducing the IR energy thereof. The turning vane arrangement (30) includes at least one turning vane (50) situated in the duct (32) and is characterized by a trailing edge (50.sub.TE) that defines an exhaust nozzle (52). The exhaust nozzle (52) is adapted for being disposed in fluid communication with a flow source (54) which produces the pressurized secondary flow PSF. The exhaust nozzle may include a plurality of adjoined lobes (60.sub.R, 60.sub.A or 60.sub.P) for rapidly and thoroughly admixing the primary and pressured secondary flows PF, PSF.

Abstract: A duct fixed to a vehicle propelled through an ambient fluid medium is innally provided with spaced channel passages from which the fluid medium is ejected under pressure tangentially of local duct surfaces through Coanda affected slots at the trailing edge of the duct from which only the ejection of the fluid medium occurs. The supply of the pressurized fluid medium under selective control is limited to different angular segments of the channel passages in order to modify the flow stream through the duct so as to perform certain functions such as thrust control and steerage control effects enhancing vehicle maneuverability.

Abstract: An exhaust nozzle for a gas turbine engine where pitch and yaw vectoring is achieved using fluidics by injecting high pressure air into the exhaust stream. The injected high pressure air deflects the exhaust stream, vectoring the aircraft accordingly. The injected high pressure air for vectoring is selectively injected. A pivoting manifold 44 allows for injection of either ram air (52) or vectoring high pressure air (56) into the exhaust flow.

Abstract: A method and apparatus for controlling the throat area, expansion ratio and thrust vector of an aircraft turbine engine exhaust nozzle, includes means, such as deflectors and/or injected air, for producing and controlling regions of locally separated flow. The exhaust nozzle also allows control of the thrust vector angle defined by the gas exiting the nozzle to provide increased directional control of the aircraft.

Abstract: An after-burning turbo-fan engine system (10) includes turbo-fan engine (14) for producing exhaust gas (20), and jet thrust therefrom. After-burning chamber (16) receives turbo-fan exhaust and injects a controllable amount of fuel into turbo-fan exhaust to cause after burning of the turbo-fan exhaust to produce after-burned exhaust (20). Nozzle (18) associates with after-burning chamber (16) for receiving after-burned exhaust (20). Nozzle (18) has a fixed geometry. Nozzle (18) flow coefficient control mechanism (22 and 24) controls the approach of after-burned exhaust (20) through nozzle (18) to change the nozzle (18) flow coefficient. This controls the volumetric flow rate of after-burned exhaust (20) through the nozzle (18).

Abstract: Nozzles for discharging high velocity air are positioned on either side of an aircraft engine exhaust duct exit. High velocity air is discharged tangential to an outer surface of each nozzle and follows the contour thereof. The nozzles are positioned so that when the high velocity air separates from the outer surface of the respective nozzles, it flows outwardly relative to the aircraft's fuselage structure. The high velocity air from the nozzle closest to the aircraft's fuselage structure impinges on the exhaust gas stream, deflecting the stream away from the fuselage. The high velocity air from the nozzle furthest from the fuselage creates a low pressure area which deflects the exhaust gas stream away from the fuselage. In combination, the two nozzles deflect the exhaust gas stream away from the fuselage to a significant degree.

Abstract: The present invention relates to a strut assembly which may be selectively projected and retracted with respect to associated structure of a solid or liquid propellant rocket of missile, with such structure comprising one of, for example, a nozzle thereof or the body thereof. The invention struct may be projected either internally of the nozzle or externally of the missile body, as desired. Projection and retraction of the strut are accomplished through use of fluid actuation means, the fluid mode of means of which comprises coolant which is also directed in various locations on the inventive strut assembly to cool the strut assembly and to protect it and adjacent structures from particulate impaction, heat and thermal damage, etc.

Abstract: A hydraulic jet propelling apparatus has an inlet opening at one end and an outlet opening at the opposite end for a fluid passing therethrough. Between the inlet and outlet openings, a centrifugal accelerator member defines a divergent annular passageway for the fluid, with a frusto-conical external sleeve and a frusto-conical central vaned rotor member keyed together on a driving shaft. A stationary cup shaped wall projects from an outlet of the accelerator member, an internal face of which, seen in axial section, having a curvature variable from the outlet of the accelerator member to define at an outer end thereof a straight direction substantially axial to said internal face. The internal face being provided with a plurality of vanes for guiding and progressively changing direction of the accelerated fluid from a diverging and rotative trajectory when leaving the accelerator member to a trajectory parallel to the straight direction when leaving the wall.

Abstract: Apparatus for modulating the thrust vector of a rocket motor by injecting t gas into the divergent section of the rocket nozzle and modulating injection of the hot gas by varying the flow from a solid propellant gas generator by controlling its flow rate with a vortex throttling valve arrangement.

Abstract: A high pressure power source for receipt on board a missile. The power source is passive until after the missile has been fired. The power source is in the form of a high pressure gas which is received in a combustion chamber of a pressure bottle. The energy from the gas is converted to electrical, pneumatic or hydraulic power as required on the missile thereby eliminating the need of electrical batteries on board the missile.

Abstract: This invention concerns the piloting of propulsion systems for solid rocket propellant missiles.It consists of a piloting assembly in which the front section of a thrust nozzle integrated in the body of a propulsion system is pierced with at least three orifices which can be covered by means of valves and which are designed to provide direct injection into the thrust nozzle of hot gases from the powder block of the propulsion system to pilot the missile. An auxiliary block of "cold" powder is fitted around the nozzle on the input side of the orifices.The application is for piloting of ballistic missiles.

Abstract: A water-jet drive mechanism for driving of watercraft and including an elbow located in the hull of the watercraft and containing a pump or the like by which water is sucked in at the bottom of the watercraft through a suction passage and is ejected downwardly through an ejector passage. A gas is supplied to the water jet at the ejector passage.

Abstract: For controlling the attitude of a rocket including a rocket fuselage and a thrust nozzle integrally attached to the rear end of the rocket fuselage and having a nozzle throat located longitudinally intermediate thereof, a rocket attitude control apparatus comprises a plurality of thrust vector control units disposed at the outer periphery of the thrust nozzle between the nozzle throat and the rear end of the thrust nozzle in circumferentially equiangularly spaced relationship to each other, and each including a fluid injecting nozzle projectable and retractable into and out of the thrust nozzle and having an injecting bore therein, wherein said fluid injecting nozzle is permitted to project into the thrust nozzle while being cooled by fluid injected from the injecting bore into the thrust nozzle. The projection of the fluid injecting nozzle causes a stream of combustion gas passing through the thrust nozzle to be partially disturbed for controlling the attitude of the rocket.

Abstract: A jet engine comprises a housing which defines a combustion chamber section, a narrowed diameter neck thrust nozzle section and an outwardly diverging portion which includes a part of the thrust nozzle section end in an extension outwardly from the thrust nozzle section. The outer end of the extension or the continuation of the thrust nozzle ends in a discharge opening for thrust gases and the vector of these thrust gases is controlled by control fluid which is selectively added into the divergent section at a location along its length such that it will be added in a selected narrow thrust jet zone in which the internal pressure on the nozzle edge of the thrust jet attains the ambient pressure or even drops below it.

Abstract: A filter-cooler is provided for use with a rocket motor to cool and filter ighly aluminized hot gases from around 5800.degree. F. to around 2000.degree. F. The cooled and filtered gases are then usable, for example, in thrust vector control or other mechanisms associated with the rocket. In the present system the cooler section is placed ahead of the filter section so that alumina present in the gases begins to solidify ahead of the filter and thus may be more efficiently removed. The filter section comprises a plurality of graphite vortex producing geometries. The vortex direction of the baffles is reversed in succeeding baffles so that the net effect of the high rotational velocities of each vortex is used to separate the particulate contaminates from the gas stream by momentum.

Abstract: This invention is a group of 8-beta, 12-alpha-PG.sub.2 (prostaglandin-type) analogs having variable chain length, or methyl or phenyl substitution in the hydroxy-substituted side-chain, and processes for making them. These compounds are useful for a variety of pharmacological purposes, including anti-ulcer, inhibition of platelet aggregation, increase of nasal patency, and labor inducement at term.

Abstract: An overexpanded thrust vector control nozzle has a converging extension section downstream of the jet exhaust stream deflecting section for defining a restriction or orifice. The orifice functions to elevate the exhaust pressure in the deflection section, which permits a reduction in nozzle size for operation at high and low altitudes. The orifice, which isolates the deflection section from outside ambient pressure, receives a deflected exhaust stream after it is turned in the direction opposite to that of the original stream deflection by the wall of the converging section. Control ports, which are adapted to communicate with a source of pressurized control fluid, are provided in the deflection section to position the exhaust stream.

Abstract: Aerodynamic surfaces are fixedly mounted in the diverging section of a rot nozzle and a fluid is injected through the fixed jet vane to flow about the vane. Fluid flow is controlled to make it appear that from upstream the jet vane is at an angle of attack. This deflects the main flow and produces side force. A minimum of injection flow enables high side forces to be obtained for missile control.

Abstract: Gas turbine engine power plants sometimes are provided with thrust reversing equipment, the outlet of which is close to the power plant air intake. Such arrangements create re-ingestion problems. There is described herein an apparatus which either maintains the upstream portion of reversed flow in a given position with respect to the plane of the air intake or alternatively destroys the reversed flow entirely.

Abstract: A control system for providing directional control to a missile during bo phase operation. The kinetic energy of the propellant exhaust gases is utilized to pump a vacuum in an annulus located around the rocket nozzle. To provide a reaction force to guide the missile, a closure is removed from one of several reaction nozzles located around the periphery of the aft end of the rocket and in communication with the vacuum chamber and the atmosphere. A thrust is generated by the air flowing through the reaction nozzle, into the vacuum chamber responsive to opening predetermined reaction nozzles.

Abstract: Fluidic switches characterized by a first divergent region immediately downstream of a nozzle throat and a second downstream region of reduced divergence which will be contacted by a deflected jet. The switches of the present invention include control ports which are normally open to the ambient atmosphere, the ports being located at or near the free stream separation point, and the transition from the first divergent section to the second section is located at or near the attached stream separation point. The switches of the present invention may assume a circular cross-section and when so doing will be provided with a plurality of ribs on the internal surface for preventing circumferential flow about a deflected jet.