Abstract: A strain measurement device includes a reference material, and a displacement sensor configured to detect relative changes in distance between the sensor and the reference material. At least one of the displacement sensor and the reference material is coupled with a pre-cured elastomeric material. The displacement sensor generates a data signal to a processor that is configured to determine a strain of another elastomeric material based at least in part on the data signal received from the sensor. A displacement sensor and a reference material may be positioned within an elastomeric material within a casing of a solid rocket motor for determining strain experienced by the elastomeric material, such as the propellant of the solid rocket motor. A method includes installing a sensor of an elastomeric material. Another method includes determining strain of an elastomeric material of a solid rocket motor.

Abstract: A hybrid rocket motor is manufactured by photopolymerizing the solid fuel grain in a stereolithography method, wherein fuel grains in a plastic matrix are deposited in layers for building a solid fuel rocket body in three dimensions for improved performance and for a compact design, the hybrid rocket motor including radial channels for defining a desired burn profile including the oxidizer to fuel burn ratio.

Abstract: The present invention relates to a solid rocket motor control system that uses state of the art electrically ignited, extinguishable and throttleable propellants. In embodiments, the control system is implemented by a software algorithm embedded in missile flight controllers and contains all control system elements and control compensation that performs thrust control for many defense and space rocket motor applications, for example.

Abstract: An end burning type gas generator includes a hollow cylindrical combustor, a propellant that is densely charged in the combustor concentrically and in a plurality of layers without leaving a large void in an internal cross section and that has the same sectional area in each of the layers, and an interstructure separating the plurality of layers from each other and connecting only at a turning portion at an axial end. The propellant end burns sequentially, progressing via the turning portion, beginning with an outermost layer or an innermost layer. The turning portion includes an end outer plate continuously surrounding ends of inner propellant and outer propellant, and a burning rate increasing member provided along the inner surface of the end outer plate to increase the burning rate of the propellant. The end outer plate is shaped such that the burning area at the turning portion remains substantially constant.

Abstract: The present invention describes a hybrid rocket motor that includes a first solid reactant and at least one thrust nozzle and at least one moveable combustion control member within the hybrid rocket motor that restricts the contact of a first fluid reactant in the combustion chamber. In this way, it is then possible to regulate the exposure of the solid reactant to the fluid reactant and thus control thrust.

Abstract: A multi-pulse rocket is equipped with a thermally insulating barrier that serves as a rupture disk or a movable plug or plate separating staged propellant grains. When a rupture disk it used, the disk can be equipped with a pyrotechnic actuator to weaken the disk upon command, enabling the propellant grain on the fore side of the disk to burst the disk at a relatively low pressure when ignition of the propellant grain is needed for additional thrust. Rupturing of the disk can also be controlled by attitude maneuvering ports on the fore side of the barrier whose open or closed conditions are controlled by independently operable closures. When a movable plug is used, the plug is movable between a closed position separating rocket chamber into subchambers isolating the propellant grains from each other and an open position allowing the flow of combustion gas between the two subchambers to achieve additional axial thrust.

Abstract: A motor includes an annular solid fuel between an inside diameter and an outside diameter. The solid fuel has a series of radial notches that define segments of fuel between them. The notches allow for faster burning of the fuel, while still allowing structural integrity of the fuel segments to be retained during the burning process.

Abstract: There is disclosed a solid fuel rocket motor which may include a case and a nozzle coupled to the case. A plurality of fuel pellets may be disposed within the case. An igniter may be disposed to ignite at least a portion of the fuel pellets. A pellet retainer may be positioned within the case to retain the plurality of fuel pellets within the case. The pellet retainer may be perforated to allow exhaust gases to flow from the ignited fuel pellets to the nozzle while preventing unburned fuel pellets from being expelled through the nozzle.

Abstract: A solid propellant thrust control system, method, and apparatus for controlling combustion of solid propellants in an opposing grains solid propellant rocket engine (OGRE) is provided. In particular, an opposing grains rocket engine and propulsion system is provided, in which actuator means connected are connected to solid propellant grains disposed in the pressure vessel of the engine. The actuator means are operable to selectively move the solid propellant grains together or apart relative to one another, such that the burning ends of the solid propellant grains decrease or increase relative to one another. This action controls the rate of combustion of the solid propellant grains by varying spacing distance between the burning ends of the solid propellant grains, and enables extinguishment and reignition of the OGRE. Further, a method is providing for controlling the burn rate of solid propellant grains undergoing combustion in an opposing grains rocket engine.

Abstract: Systems and methods of controlling solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and propellant gas flow rate, rely on pulse width modulation of a control valve duty cycle. A control valve that is movable between a closed position and a full-open position is disposed downstream of, and in fluid communication with, a solid propellant gas generator. The solid propellant in the solid propellant gas generator is ignited, to thereby generate propellant gas. The control valve is moved between the closed position and the full-open position at an operating frequency and with a valve duty cycle. The valve duty cycle is the ratio of a time the control valve is in the full-open position to a time it takes the valve to complete one movement cycle at the operating frequency. The valve duty cycle is controlled to attain a desired solid propellant burn rate, propellant gas pressure, propellant gas pressure pulse shape, and/or propellant gas flow rate.

Abstract: Motor designs that enable the control of ignition, combustion burn rate, extinguishment, and reignition of solid propellants by precise application of electrical power are provided. Design of such motors, including choice of electrode materials, the form of electric power, and exemplary facial-extent electrode and axial-extent electrode configurations are also provided.

Abstract: An apparatus and method to control ignition, combustion burn rate, extinguishment, and reignition of solid propellant is provided.

Abstract: A side thruster of a flying object of the invention is provided with a solid gas generating agent, a nozzle arranged in a direction perpendicular to an axis of the flying object, a tube body for flowing the gas generated in accordance with a combustion of the solid gas generating agent to the nozzle, the solid gas generating agent being arranged around the tube body, a plurality of orifices provided on a peripheral wall of the tube body along its circumferential direction and capable of injecting the gas generated in accordance with the combustion of the solid gas generating agent toward an axis of the tube body, and a valve provided within the tube body and rotatable around the axis of the tube body so as to open and close the plurality of orifices.

Abstract: A solid divert and attitude control system for use on a kill vehicle or missile that includes a solid propellant gas generator, a control valve for modulating the effective exit area of the solid propellant gas generator, and a controller for calculating the integral of the burn rate of the solid propellant and for controlling operation of the control valve in response to the calculation of the integral of the burn rate of the solid propellant. The solid propellant gas generator has a solid propellant with a non-constant burn area and a grain profile which includes a first boost phase grain, an intermediate sustain phase grain, and a second boost phase. The controller controls operation of the control valve during a coast phase of the mission so that the second boost phase grain does not ignite until the beginning of the terminal phase of the mission.

Abstract: Solid-propellant rocket engine with a combustion chamber, with a thrust nozzle, with one or more grains in the combustion chamber, at least one of which comprises two parts, namely, an inner burner part and an outer burner part, wherein the latter is connected to a support structure, has a cross section changing over its length, is accommodated in the hollow space of identical shape of the inner burner part, and is axially displaceable. A thermal barrier is arranged at the nozzle-side end of the two-part grain. The outer burner part is fitted gap-free into the inner burner part. The contour formed by the contact surfaces of the two grain parts encloses a cross section expanding from the front to the rear. The outer burner part can be moved from a front, inactive position, which is positive-locking with the inner burner part, into a rear, active position with a defined flow gap between the contact surfaces of the grains.

Abstract: A controlled fuel flow system from variable flow solid fuel gas generator to secondary combustor is provided, which system includes a hot gas valve which is operated by an electronic controller. The valve is mounted between the gas generator and the secondary combustor, which valve has in the throat thereof a pivotable blade that moves to relatively blocking and unblocking positions therein and a nozzle downstream of such blade which communicates with the secondary combustor. A first pressure transducer is mounted in the gas generator and a second pressure transducer is mounted downstream of such blade near or in the nozzle. Actuator means are also provided to pivot such blade between relatively blocked and unblocked settings as guided by electronic gas flow controller logic. The secondary combustor has one or more air ducts therein and the above system is mounted, e.g., in a variable flow ducted rocket.

Abstract: The device comprises, at the outlet of the combustion chamber 3, a movable assembly comprising a hollow closure member 26 constituting a needle 27, a thermal shield 22 and an annular piston 14. This closure member acts as a piston under the effect of the pressure of the gases and drives the annular piston so as to expel a hydraulic fluid from a jack chamber 33 having an adjustable outlet 34. The closure member 26 defines a chamber 28 for calming the gases.

Abstract: A velocity controller for a ramjet missile, having a supersonic inlet proximate the peripheral skin thereof for admitting air to a combustion zone of a ramjet engine, is comprised of a variable pitch cover disposed in pivotable engagement within the inlet and an actuator in operative engagement with the cover for adjustably positioning same over an angular range and thereby modulating airflow for the purpose of controlling flight characteristics and, principally, velocity of the missile. A sensing system is provided for detecting a dynamic flight parameter indicative of velocity of the missile and generating an output characteristic thereof for controlling the actuator and, in turn, the pitch of the cover. Methods for improving the flight performance of both solid fuel ramjet missiles and ducted rocket missiles are also disclosed herein.

Abstract: The utilization of metal hydride and an acidic reagent for the accelerating of masses, and propulsion devices for applying such materials. The materials are of a saline complex metal hydride and acidic reagent which is sprayed onto the metal hydride which is offered in a lumpy consistency, for the pulse-like generation of expanding reaction gas bubbles in a constructive or dynamically, yieldably dammed chamber. Also provided is a propulsion device for the application of the above materials, with a support for a saline complex metal hydride of a lumpy consistency in a yieldably dammed expansion chamber in which an injection nozzle for the reagent is directed towards a surface portion of the metal hydride.

Abstract: A solid fuel rocket burn rate control method wherein grain preheating energy is supplied to the grain region just ahead of the regressing burn face by optical conductors such as fiber optic filaments that are buried in the grain and dispersed across the grain cross section. The optical conductors receive optical energy from one of several types of electrical-to-optical energy transducers such as a semiconductor laser or an incandescent source; the optical conductors promote coning burn action at the burn face of the rocket and allow control of the burn rate by electrically modified optical signals.

Abstract: A solid fuel rocket burn rate control apparatus wherein grain preheating energy is supplied to the grain region just ahead of the regressing burn face by optical conductors such as fiber optic filaments that are buried in the grain and dispersed across the grain cross section. The optical conductors receive optical energy from one of several types of electrical-to-optical energy transducers such as a semiconductor laser or an incandescent source; the optical conductors promote coning burn action at the burn face of the rocket and allow control of the burn rate by electrically modified optical signals.

Abstract: A method for steering solid propellant ballistic vehicles during powered flight which eliminates the requirement for cutoff control by allowing simultaneous fuel depletion and velocity-to-be-gained, V.sub.G, nulling. The vehicle booster is steered along a velocity trajectory of length equal to the remaining velocity capability, V.sub.CAP, which results in a fuel-inefficient trajectory. The trajectory is divided basically into three phases--an exit phase, a fuel-depletion guidance (FDG) phase and a short phase of constant attitude thrusting just prior to final stage burnout. For the exit phase the launch azimuth and the pitch-over magnitude can be varied from their usual fuel-efficient values. During fuel-depletion guidance the additional degree of freedom is the angle, .theta., between V.sub.G and the desired thrust direction, U.sub.S, where: ##EQU1## a.sub.T being the sensed acceleration vector.

Abstract: A solid propellant rocket motor of gas generator of the type having retractable filaments embedded in the propellant for the purpose of controlling the overall thrust or mass flow output is combined with a channeled separator located between the propellant and its casing with the channels open to the combustion chamber. This results in distribution of combustion chamber pressure and temperatures over the exterior of the propellant grain so that distortion of the grain will not occur sufficiently to cause binding or sticking of the filament in their holes in the propellant nor will the holes be enlarged to cause uncontrolled burning between the filament and the hole.

Abstract: Thrust termination of a solid propellant rocket motor is obtained by connecting the nozzle to the combustion chamber pressure vessel by means of explosive and extrusable bolts. When it is desired to obtain thrust termination, the explosive bolts are blown leaving the nozzle connected to the pressure vessel by the extrusable bolts. The extrusable bolts have sufficient strength to retain the nozzle assembly; however, upon elimination of the strength of the explosive bolts, the nozzle is permitted to separate from the pressure vessel and move to the rear by drawing the extrusable bolts through dies. The bolts deform to absorb the energy of the nozzle and the jolt of forward acceleration which would normally be obtained if the nozzle were ejected, is reduced or eliminated.