Abstract: The jet engine comprising a ramjet air path extending from an intake, into a combustion chamber, and out an exhaust nozzle, a fuel inlet leading into the combustion chamber, an oxidizer inlet leading into the combustion chamber and a partition being operable to selectively close the ramjet air path upstream of the combustion chamber to allow operation of the jet engine in rocket mode and open the ramjet air path to allow operation of the jet engine in ramjet mode.

Abstract: A turbine pump assembly has a turbine, a centrifugal pump, and a passive electrical speed control system. The turbine has a peak efficiency at a first speed that is lower than a second speed at which the centrifugal pump is operating at a peak power requirement. A rocket thrust vector control system is also disclosed.

Abstract: A selectable ramjet propulsion system for propelling a rocket or missile includes a gas generator adjacent a booster. A frangible diaphragm is disposed between the gas generator and the booster. The booster and fuel gas generator can be operated in normal sequence, or operated at the same time in order to increase the thrust produced for short-range missions. A logic circuit contained on the rocket or missile determines a time to rupture the frangible diaphragm based on whether or not the distance to the target exceeds a threshold distance.

Abstract: Systems and methods are described herein for a hybrid liquid propellant rocket engine. In an embodiment, the engine includes a first pump powered by a first turbine, a second pump powered by a second turbine, and a gas generator. An output of the gas generator is connected to the first turbine and the second turbine. The engine further includes a third pump powered by a third turbine, a fourth pump powered by a fourth turbine, and a nozzle having an expander cycle in a wall and a combustion chamber. An output of the third pump is connected to the expander cycle and an output of the wall is connected to the third turbine and the fourth turbine. An output of the fourth pump, an output of the third turbine, and an output of the fourth turbine are connected to the combustion chamber.

Abstract: New missile systems are provided, implementing reduced pre-deployment weight, higher impact and greater deployment flexibility, among other advantages. In some embodiments, mid-flight oxygen filtration from atmospheric air, followed by concentration, compression and/or storage in ideal oxidizer deployment locations, leads to enriched, greatly increased oxidizer load and/or far greater missile weight just prior to impact. Among additional benefits, missiles implementing the system may be far less volatile, and therefore safer, prior to deployment and the concentration of oxidizer may be more concentrated than with ambient oxygen, overcoming the limitations of current fuel/air and other thermobaric explosives.

Abstract: The method serves to calculate a start sequence or a stop sequence for an engine, said sequence comprising a plurality of cues at which predetermined actions should be performed in the control of said engine. The method comprises: a step of obtaining at least one thermodynamic condition of said engine; and a step of calculating said sequence as a function of: said at least one thermodynamic condition; dimensional parameters of said engine; and criteria for proper operation of said engine.

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 ram jet engine including a combustion chamber ending in a gas injection nozzle, a cruising propulsion unit fee ding liquid propellant into the combustion chamber, and at least one air duct for feeding air intended for combustion of the fuel in the combustion chamber. A rigid tubular element is incorporated into the ramjet engine, with an interior volume of the rigid tubular element being divided into two spaces by an intermediate transverse partition. One of the spaces houses the cruising propulsion unit while the other space houses the combustion chamber. Passages are cut into the intermediate transverse partition to allow feeding of liquid fuel into the combustion chamber, and the air duct is mounted on the tubular element to feed the combustion air through the tubular wall of the tubular element.

Abstract: An ultra-compact aerovortical swirl-enhanced combustion (ASC) system features an aerovortical swirl generator for use in rocket thrusters utilizing hypergolic or non-hypergolic propellants. The ACS thruster can be sized for diameters ranging from about 0.5 to about 2.0 inches, and producing thrust levels of approximately 5 lbf to about 250 lbf. A plurality of helicoid flow channels in the swirl generator introduces swirl into a flow stream of a first propellant within ultra-compact sized rocket thrusters. The ASC system also includes injectors for introducing a second liquid propellant into the swirling flowfield to promote rapid and efficient atomization, mixing and vigorous combustion, which, results in major improvements in combustion and propulsion performance over current rocket thrusters, but in much shorter combustor systems.

Abstract: A ramjet powered device that utilizes a novel swirl generator for rapidly and efficiently atomizing, vaporizing, as necessary, and mixing a fuel into an oxidant. The swirl generator converts an oxidant flow into a turbulent, three-dimensional flowfield into which the fuel is introduced. The swirl generator effects a toroidal outer recirculation zone and an inner central recirculation zone, both of which are configured in a backward-flowing manner that carries heat and combustion byproducts upstream where they are employed to continuously ignite a combustible fuel/oxidizer mixture in adjacent shear layers and stabilizes flame propagation and accelerates combustion throughout the entire combustor. The swirl generator provides smooth combustion with no instabilities and minimum total pressure losses and enables significant reductions in the L/D ratio of the combustor. Other benefits include simplicity, reliability, wide flammability limits and high combustion efficiency and thrust performance.

Abstract: A blanking-plug system for blanking off an orifice of a pipe, particularly for blanking off an orifice of a duct for introducing air into the combustion chamber of a ramjet. The blanking-plug system (16) comprises at least one blanking plug (15) comprising a glass plate (18) which is able to completely blank off said orifice (8) of the pipe (7); and at least one device (17) for destroying said glass plate (18) of said blanking plug (15). Said blanking plug (15) additionally comprises a plurality of elastomer protective elements (21) which are fixed on at least one first face (22) of said glass plate (18) which is likely to be subject to attack, these elements being separated from one another by a predetermined maximum distance at most, and completely covering said first face (22).

Abstract: A blanking-plug system for blanking off an orifice of a pipe, particularly for blanking off an orifice of a duct for introducing air into the combustion chamber of a ramjet.

Abstract: In an integral rocket-ramjet engine the air intake ports are covered with a port cover made of a laminate of Pd and Al. This port cover is rapidly consumed in the brief period between the end of the rocket mode and commencement of the ramjet mode to allow ingress of ram air for the ramjet engine.

Abstract: Blanking-cap system for an orifice of a duct, in particular for an orifice of an air-intake duct into the combustion chamber of a ramjet. The blanking-cap system (16) includes a glass blanking cap (15) and a destruction device (17) which comprises a projectile (18) capable of destroying the blanking cap (15), and a controllable projection means (19), which is capable of projecting the said projectile (18) and which is arranged outside the conduit (7) while being oriented in such a way as to be able to project the said projectile (18) onto the said blanking cap (15).

Abstract: This multi-stage rocket motor assembly includes, a one main motor subassembly and a launch motor subassembly. The flight igniter of the main motor subassembly is integrated with the launch motor subassembly to permit the flight igniter to be jettisoned essentially in unison with the launch motor subassembly.

Abstract: A solid fuel propulsion system for a ran jet rocket having a combustion chamber (B) surrounded by a tubular casing (2) and a gas generator (G) disposed upstream from the combustion chamber (B) and surrounded by a tubular casing (1) for generation of a combustible gas from a solid fuel, disposed between the gas generator (G) and the combustion chamber (B) is a gas-stream regulator unit (R) for regulation of the flow of the combustible gas from the gas generator (G) to the combustion chamber (B). The propulsion system has a middle section (4) in which the gas stream regulator unit (R) is housed. The middle section is connected in a load-bearing manner with the combustion chamber casing (2) and the gas generator casing (1) and is provided with a first pressure head (8) sealing the gas generator (G), and a second pressure head (9) sealing the combustion chamber (B). Disposed between the pressure heads (8, 9) is a base unit which houses the gas stream regulator unit (R) and braces the pressure heads (8, 9).

Abstract: The invention is an airframe which includes a vehicle (12) having a solid propellant rocket engine (14) and a ramjet or scramjet engine (16); a thrust plug (18) extending from an end (20) of the vehicle which directs combustion gases (23 and 64) produced by the solid propellant rocket engine or ramjet/scramjet engine to produce forward thrust; a longitudinal passage (38) extending from the end of the vehicle to an opening (30) forward of the end which receives external air directed by forward movement of the vehicle and in which solid propellant (32) of the solid propellant rocket engine is located, and wherein during rocket operation solid propellant is combusted to produce the combustion gases in the longitudinal passage which are conveyed by the longitudinal passage into contact with the thrust plug and during ramjet/scramjet operation the longitudinal passage is open to flow of external air after operation of the solid propellant rocket engine is completed and which supports mixing and combustion of the a

Abstract: The present invention relates to a hybrid engine capable of employing a ramjet mode and a super ramjet mode and comprising a main engine system (3, 4, 5, 8) which burns fuel and which comprises at least one air inlet (3) for an air stream (F), a combustion chamber (8) and a jetpipe (4). According to the invention, said engine (1) additionally comprises first means (10) for storing, in ramjet mode, air diverted from the air stream (F), and second means (11) for burning, in super ramjet mode, some of a fuel of the engine (1) with the air stored by the first means (10), before injecting it into the main engine system (3, 4, 5, 8).

Abstract: The present invention relates to a shut-off system for an orifice for introducing combustion air into a ramjet with a consumable auxiliary motor (14).According to the invention, this shut-off system comprises:an elastic system (28, 29, 31, 32) connected to the shut-off flap (20A, 20B); anda retaining element (33) for keeping the said elastic system in the said tense state during the initial phase of rocket operation, the said retaining element being sensitive to the hot gases emitted by the said consumable auxiliary motor (14) so that on completion of combustion thereof, the said retaining element (33) releases the said elastic system which spontaneously changes from its tense state to its relaxed state, bringing the said flap from its shut-off position to its open position.

Abstract: A missile adapted for flight at hypersonic velocities includes an engine operable in rocket and ramjet modes of operation, the engine having an inlet opening, a fuel combustion chamber in the engine housing a boost fuel and a cruise fuel, an axially movable plug located at the engine inlet opening for opening and closing the inlet opening, and a mechanism, coupled with the plug and the engine, for switching between the two modes of operation of the engine during flight of the missile. In this way, when the missile reaches a target location in its flight trajectory, the plug can be moved to close the inlet opening and shut down the ramjet operation, while also minimizing the missile radar cross section properties. The switching mechanism includes sensors for determining flight parameters and a computer for processing the flight parameters to determine when to move the plug.

Abstract: A pulse detonation rocket engine, having at least two detonation chambers. The rocket propelled vehicle includes at least one fuel delivery system in fluid communication with each of the at least two detonation chambers, and at least one oxidant delivery system in fluid communication with the detonation chambers, along with fast-acting valves to inject fuel and oxidant controlledly into the chambers. An ignitor in each of the detonation chambers intermittently initiates detonation of a fuel and oxidant mixture in the chamber, in a controlled cycle, to provide motive force. Also provided is a combined cycle engine, able to operate in air breathing mode, oxidant augmented mode, and as a rocket engine. The combined cycle engine includes at least one detonation chamber, and may include a plurality of such chambers. The invention further provides methods of intermittently detonating sequentially created fuel and oxygen mixtures in these engines, and methods of using these engines.

Abstract: The present invention discloses a composite material closure as a port cover for a rocket-ramjet engine and a method of transitioning between rocket booster operation and ramjet operation. The composite material closure is made to be coextensive with an insulating lining of the casing of the rocket-ramjet engine. Pyrotechnic charge devices are associated with the composite material closure and ignited to cut there-through and form an opening in the closure for air flow during ramjet operation.

Abstract: A multi-pulse or multi-stage canister loaded solid propellant rocket motor. The canisters are prepared separately and loaded with solid propellant whereby the scrape rate may be reduced after which they are installed in a monolithic case which affords stiffness continuity over the length of the rocket motor to prevent guidance system upsetting free play. In order to reduce the complexity of installation of the membrane seal assembly for each pulse, the bulkhead therefor is manufactured integral with the case of the respective canister and becomes the forward closure thereof. For the multi-stage rocket motor, the monolithic case may be stepped and tapered, and the stages severable therefrom as they burn out by primer cord.

Abstract: A pyrotechnic ignition train for use in dual nozzle rocket-ramjet missiles herein a donor device and a receptor charge in the train allow transfer of a pyrotechnic impulse across an air gap between an outer nozzle and an inner concentric nozzle to allow ignition of the rocket motor. Upon consumption of the rocket fuel the inner nozzle is jettisoned, and the ramjet motor fired. Absence of physical connection of the nozzles at the ignition train allows easy separation of the inner nozzle from the missile.

Abstract: The invention disclosed is a solid propellant rocket motor, capable of providing two separate propulsive impulses to a missile. The rocket motor is connected at one end to the missile body, the other end including an exhaust nozzle. The rocket motor comprises two stages connected by an interstage bulkhead. The bulkhead includes a port opening which is closed by a frangible cover which prevents the second stage from igniting during burning of the first stage, but breaks up into harmless fragments during firing of the second stage.

Abstract: A multi-pulse or multi-stage canister loaded solid propellant rocket motor. The canisters are prepared separately and loaded with solid propellant whereby the scrap rate may be reduced after which they are installed in a monolithic case which affords stiffness continuity over the length of the rocket motor to prevent guidance system upsetting free play. In order to reduce the complexity of installation of the membrane seal assembly for each pulse, the bulkhead therefor is manufactured integral with the case of the respective canister and becomes the forward closure thereof. For the multi-stage rocket motor, the monolithic case may be stepped and tapered, and the stages severable therefrom as they burn out by primer cord.

Abstract: A solid propellant rocket motor propellant grain configuration having a center bore of a varying diameter, ballistic slots, a stress/ballistic groove, and a burn inhibitor band for withstanding service motor operating environments and providing the required ballistic profile.

Abstract: A solid rocket motor grain construction is disclosed for use in a ramjet to improve performance. The solid fuel burn enhancer adds oxidizer for burn rate control. Two constructions are disclosed. The inlaid approach places cores of solid low-oxidizer fuel within solid unoxidized fuel and the integrated approach provides shells of low oxidizer around the solid unoxidized fuel. The low-oxidizer fuel has a faster burn rate and smaller burn area requirements than the unoxidized fuel. During initial ramjet operation, when the burn area is constrained, the low-oxidizer fuel burns initially with its smaller burn area requirements. Subsequently, the conventional unoxidized fuel burns when larger burn areas are available.

Abstract: An integrated rocket ramjet in which initial acceleration of the rocket is achieved by a rocket fuel propellant and subsequent flight is sustained by a ramjet engine receiving air via a supersonic diffuser from an inlet disposed at the middle of the rocket. The flame for the ramjet combustion system is incorporated into the diffuser so that diffusion and combustion take place simultaneously. The combustion process continues within the space previously occupied by the rocket fuel propellant and the walls bounding this space are the outer casing of the rocket and are directly cooled by free stream air flowing past the rocket.

Abstract: A rocket engine according to the invention has a body generally extending along an axis and adapted to move through the air in a predetermined axial back-to-front direction. This body is formed with axially spaced front and rear fuel-containing combustion chambers. A nozzle centered on the axis and extending axially backward in the body from the front chamber has a rear nozzle end directed axially backward so that gases generated in the front chamber by combustion of the fuel therein form a fan-shaped jet extending axially backward from the rear nozzle end.

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: A missile bulkhead rupture disc assembly for disposition between adjacent solid propellant stages which are ignited in sequential order has a central aperture which receives an igniter for detonating the first stage of propellant. In preferred forms, the rupture disc comprises a pre-bulged frangible element having a first burst pattern with a circular line of weakness circumscribing a central portion of the element surrounding the igniter, and the element also has a second burst pattern comprising a plurality of lines of weakness which extend in a radial direction outwardly from the circular line of weakness of the first burst pattern, and which are spaced from each other to define generally sector-shaped petals or segments of the frangible element.

Abstract: An integral rocket and ramjet engine (16) comprises, in flow series, an intake duct (18) for aerodynamically compressing air, a port cover (30), a combustion chamber (20) and a propelling nozzle (22). The port cover (30) in a first position prevents air from entering the combustion chamber (20) and enables a rocket charge (24) to be burnt in the combustion chamber (20) thereby accelerating the engine to sufficient velocity for ramjet operation. The port cover (30) is movable axially to a second position when the rocket charge (24) is spent to allow compressed air into the combustion chamber (20) and the engine (16) to operate as a ramjet. In the second position the port cover (30) provides a quiet zone (52) which is substantially shielded from the main flow of air into the combustion chamber. Combustion equipment such as primary and main fuel manifolds (42,46) and flame gutters (50) are also movable with the port cover (30).

Abstract: A missile drive is provided comprising a single combustion chamber shared by a first, acceleration stage and a second, ramjet cruising stage, said chamber housing the solid propellant to be consumed during acceleration. At least one air inlet is opened at the end of the acceleration stage to allow an air flow to be introduced into the combustion chamber. Moreover, at least one additional exhaust outlet forming an additional nozzle is provided in the back of the combustion chamber and means are provided to seal off said additional exhaust outlet or outlets throughout the initial acceleration stage, such that said additional exhaust outlet or outlets contribute, together with the converging-diverging nozzle, to ejecting the exhaust gas during the ramjet cruising stage.

Abstract: In a power control for a turbo-generator, the acceleration power of the difference between the mechanical power supplied to the turbo-generator set and the power delivered by it is determined and, if the supplied power preponderates, a correction signal directed toward closing control valves is fed to the power control by means of a valve positioning controller, Thereby, the power control need not be disconnected in the event of disturbances such as load shedding and network short circuit, but can practically remain engaged continuously.

Abstract: The metallic wall of the combustion chamber of a combined rocket-ramjet engine is lined with solid ramjet fuel overlaid with rocket fuel. After the consumption of the rocket fuel in the boost portion of the flight the solid ramjet fuel burns and ablates protecting the metallic combustion chamber wall from high temperatures during the cruise phase of the missile flight.

Abstract: A pyrotechnic delay system for controlling the time of events from initian of an initial ignition impulse to ignition of a secondary output.

Abstract: An improved combination of a cover for closing the air inlet openings into he combustion chamber of a jet rocket engine and strike device for destroying the cover is disclosed. In accordance with the improved arrangement, the cover is composed of a material which is internally prestressed and a device is provided for retaining the strike device in a prestrike position relative to the cover which is operative to release the strike device upon the reaching of a predetermined magnitude of the pressure acting on the strike device.

Abstract: A thrust nozzle for a reaction engine having a combustion chamber with a sustained flight discharge nozzle defined therein for the passage of thrust gases and a sustained flight thrust arrangement connected to the combustion chamber for supplying thrust gases through the combustion chamber, the thrust nozzle comprising a starting nozzle formed on an interior wall of the combustion chamber radially inwardly of the sustained flight discharge nozzle made of a plurality of layers of ablatable material which are ablated by the passing thrust gases to discharge vaporized or fragmentary parts of the layers of ablatable material.

Abstract: A ramjet engine for aircraft incorporating means for automatically changing from a subsonic combustion ramjet geometry to a supersonic combustion ramjet geometry. This is accomplished by utilizing solid fuel for the ramjet during the subsonic combustion mode and tailoring the solid fuel such that, as it burns away, the remaining internal engine geometry changes to the internal engine geometry for supersonic combustion. At the desired time, liquid fuel can be injected into the combustion chamber and the ramjet will operate eventually on liquid fuel only as a supersonic combustion ramjet.

Abstract: A solid fueled ramjet engine comprising solid fuel within a combustion chamber in the form of a hollow cylinder, and a windmill at the entrance to the hollow cylinder for promoting better distribution of the air, better mixing of the air and combustion gases, and more complete combustion of the solid fuel. The windmill is turned by the incoming airflow and can rotate a generator to provide a source of electrical power for the aircraft on which the engine is used.

Abstract: A multi-burn restartable solid rocket fuel having plural layers, each independently ignitable, including a central core, and at least one outer cylindrical layer separated from the central core by a non-ignitable layer, and the method of utilization therefor including first igniting the central core in an end burned configuration, and subsequently or concurrently igniting the next adjacent layer or layers in an end burned configuration, as desired.

Abstract: A propulsion motor for an underwater vehicle such as an anti-submarine weapon. The motor includes a propulsion chamber into which water is admitted and then rapidly expelled through an exhaust nozzle, developing thrust to propel the vehicle. Gas generators are used to develop the successive hydropulses to expel the water following each filling of the motor chamber with water. In one particular embodiment of an anti-submarine weapon which is directed through the air to the vicinity of a submarine by a rocket motor, the hydropulse underwater propulsion system can use the same chamber as the rocket motor.

Abstract: A two-stage aerial vehicle which is launched and accelerated by an integral olid booster. After booster burnout and jettison of the booster nozzle, the booster chamber becomes the combustion chamber for a liquid fueled ramjet. The vehicle is designed for use with known guided missile launching systems and employs unique ram air scoop and control surface actuator structure.

Abstract: In an integral rocket-ramjet having a combustor which initially serves as a rocket combustion chamber for booster propellant, and after the booster propellant is expended serves as a ramjet combustor where fuel and air are burned, a fuel control system is described for the ramjet stage by which ram burner light-off is automatically initiated upon transition from rocket to ramjet propulsion. The fuel control regulates fuel flow to the combustor over the entire flight regime and responds to operating conditions to provide a light-off schedule, to stabilize the shock wave at the air inlet, to provide a maximum fuel-to-air ratio limit, to limit the maximum vehicle Mach number, and to prevent lean burner blowout by providing a minimum fuel-to-air ratio limit. Mach number limiting and air inlet margin limiting are performed in closed loop fashion, while the other functions are scheduled or open loop controls.

Abstract: In an integral rocket-ramjet having a combustor which initially serves as a rocket combustion chamber for booster propellant, and after the booster propellant is expended, serves as a ramjet combustor where fuel and air are burned, a fuel control system is described for the ramjet stage by which ram burner light-off is automatically initiated upon transition from rocket to ramjet propulsion. The fuel control regulates fuel flow to the combustor over the entire flight regime and responds to operating conditions to provide a light-off schedule, to stabilize the shock wave in the air inlet, to provide a maximum fuel-to-air ratio limit, to limit the maximum vehicle Mach number, and to prevent lean burner blowout by providing a minimum fuel-to-air ratio limit. Mach number limiting is performed in closed loop fashion, while the other functions are scheduled or open loop controls.

Abstract: The invention relates to a combustion chamber of a ramjet and booster rocket, which comprises an inner envelope provided with upstream lateral ports for the inlet of air and a downstream nozzle for ejecting the combustion gases, to allow functioning of the rocket at cruising speed; an outer envelope, concentric to the inner envelope and closed, with the exclusion of a single rear port for the ejection of the gases, to allow functioning of the rocket at launching speed; and separating means for separating from the combustion chamber at least a part of the outer envelope, at the end of the launching phase, in order to ensure the clearance of the upstream lateral ports of the inner envelope.

Abstract: Protectors for certain elements such as fuel injectors and flameholders used in conjunction with the ramjet phase of an integral rocket ramjet.In the preferred embodiment slip-on caps of a high temperature material such as graphite are used to protect the fuel injectors and flameholders. The caps are installed and the solid boost propellant packaged about them. Following boost burnout the initial fuel pressurization for transition to ramjet operation would remove the caps allowing operation in a ramjet mode.

Abstract: A sensor for rocket engines having at least one combustion chamber including a pressure-sensitive sensing device responsive to pressure in the combustion chamber for providing an output signal to a control device in response to the pressure in the combustion dropping below and/or rising above at least one predetermined pressure. The pressure-sensitive sensing device includes a piston arranged for displacement between first and second positions. The piston is displaceable from at least the first position to the second position in response to a first predetermined pressure value in the combustion chamber. A restoring force is effective for displacing the piston from the second position toward the first position in accordance with a second predetermined pressure value in the combustion chamber. The piston is arranged for triggering a signal in at least one of the first and second positions thereof.

Abstract: A rocket motor grain consisting of a one-piece solid base propellant divided into an aft end section having a central cavity extending throughout its length, a forward end section having a central longitudinal cavity filled with a propellant having a substantially higher burning rate than the base propellant and a central web section being defined between the forward and aft end sections. This radial end burning rocket motor grain provides a means for utilizing high energy and high density propellants which have high burning rates in a manner such that the burning rate effects can be offset to provide enhanced motor performance without the sacrifice of high energy. After initiation of the rocket motor grain, the aft end section and central web section begin burning together to provide the desired mass flow rate, followed by initiation of a higher burning rate propellant in the forward section.