Method And Plant For The Production Of A Casing For A Solid-Propellant Engine, And Casing Made According To Said Method

A casing of a solid-propellant engine comprising a core and a layer of elastomeric material, set as coating for at least part of the core to provide a thermal protection of the core itself is obtained by: inserting the core in a forming mould so as to make within the mould two annular chambers separated from one another by the core; forming a strand of elastomeric material; obtaining a defined portion of elastomeric material by cutting the strand transversely to size in an external environment; and injecting the cut portion of elastomeric material simultaneously within both of the annular chambers.

Ammunition Comprising Means for Neutralizing Its Explosive Charge

The invention relates to an ammunition comprising an explosive charge confined in a compartment and a deconfinement device capable of deconfining the explosive charge under a pressure or temperature rise within the ammunition. The invention neutralizes the ammunition on command. According to the invention, the ammunition further includes means for activating the deconfinement device, these means being controllable.

Systems and Methods for Casting Hybrid Rocket Motor Fuel Grains

Embodiments of the invention relate to systems and methods for casting hybrid rocket motor fuel grains. In one embodiment, a method for casting a rocket motor fuel grain can be provided. The method can include providing a positive image of a port made from at least one material. The method can further include disposing at least one fuel material around at least a portion of the positive image of the port. Further, the method can include removing the at least one material, wherein a negative image of the port is formed in the at least one fuel material.

Linear Detonation Wave Diverter

The presently disclosed linear detonation wave diverter provides a structure and method for quickly and controllably venting a detonation event out of the diverter without igniting working fluid upstream of a microporous barrier within the linear detonation wave diverter. Further, the detonation wave is linearly vented out of the diverter upon the failure of a burst member, which provides a low resistance path for detonation waves to exit the detonation wave diverter.

Rocket motor with means for user adjustable thrust

A rocket having a motor with a user adjustable thrust is provided. The rocket includes a main cylinder containing a rocket propellant, the propellant configured to generate gas during operation and one or more nozzles arranged to direct the gas in a first direction. A thrust adjustment device is arranged to receive a portion of the gas, the thrust adjustment device configured to change the direction of flow of at least a portion of the gas

SmallSat Hybrid Propulsion System

A hybrid propulsion system for a small satellite package consisting of a main rocket motor containing a solid propellant with multiple oxidizer tanks positioned to direct oxidizer into the rocker motor, thereby producing a desired thrust necessary for orbit insertion and/or orbit correction. Additionally, oxidizers can serve a dual function in controlling cold fuel thrusters for attitude adjustment. 1. A CubeSat propulsion system comprising:a CubeSat form factor;a main propulsion motor vessel centrally disposed within form factor and having a body with a forward end and an aft end and being formed of an outer wall and an inner wall, wherein the inner wall forms a central cavity, and wherein the inner wall of the central cavity is lined with an insulative material that forms a thermal protection layer between the inner wall of the central cavity and a solid rocket fuel disposed within the central cavity such that the fuel is positioned where the majority of the central cavity is filled with the solid rocket fuel, and wherein the aft end of the motor vessel further comprises a flight thrust nozzle having a nozzle throat section and a nozzle exit section wherein the nozzle throat section is positioned near the aft end of the body and the nozzle exit section is disposed distal to the aft end,multiple oxidizer containment vessels disposed within the form factor and dispersed around the propulsion vessel and wherein each of the oxidizer vessels are fluidly connected to the propulsion vessel, such that the oxidizer vessels deliver an oxidizer into the propulsion vessel within a predefined channel disposed within the fuel, andan ignition source mechanically connected to the propulsion vessel wherein the ignition source operates to vaporize a portion of the fuel and wherein the oxidizer delivered to the propulsion vessel interacts with the vaporized fuel to produce combustion along the length of the predefined channel and produce an exhaust thrust through the nozzle.2. The ...

ROCKET DELAY APPARATUSES, SYSTEMS AND METHODS

Delay tools, systems and methods for achieving a selection of alternative delay times, a tool of which including a body, a drill bit operable relative to the body and a knob operably connected to the drill bit, and operably disposed relative to the body for engagement of the body with a rocket motor bulkhead and the drill both relative to a delay to provide for achieving a selection of alternative delay times. 1. A delay modification tool comprising:a body;a drill bit operably disposed relative to the body; anda knob operably connected to the drill bit, and operably disposed relative to the body,wherein the body, drill bit and knob are operably configured for engagement of the body with a rocket motor bulkhead and the drill bit relative to a delay disposed within the bulkhead to provide for achieving a selection of alternative delay times.2. A delay tool according to wherein the body has at least first and second sides each being alternatively engageable with the rocket motor bulkhead claim 1 , each providing for alternative delay times.3. A delay tool according to wherein the body has at least first and second sides each having a respective first and second well of two discrete first and second well sizes to achieve two alternative delay times.4. A delay tool according to :wherein the body is substantially cylindrical andthe body has at least first and second sides each of the first and second sides having a respective first and second well of two discrete first and second well sizes each being alternatively engageable with the rocket motor bulkhead, each providing for alternative delay times.5. A delay tool according to including a spacer to alter the delay to be achieved.6. A delay tool according to wherein the spacer is removably disposable between the body and the knob and is configured so that the knob will be unable to contact the body.7. A delay tool according to further comprising internal structure configured to provide for engaging a rocket bulkhead at ...

ALTERNATIVE METHOD FOR DISMANTLING SOLID-PROPELLANT MOTORS

A method is provided for solid-propellant engines to be dismantled safely and in accordance with environmental standards having been scrapped. For each engine to be dismantled, it is mounted on a static test rig, immersed in a tank filled with water and started such that propellant is used up under the water. The soluble part of the combustion products (gases or condensates) thus remains trapped in the water in the tank while the non-soluble solid products drop to the bottom of the tank. The body of the engine emptied of its fuel in this way and rendered pyrotechnically inert is then taken apart or disassembled. Periodically, the water in the tank is withdrawn and the tank stripped of its deposits such that subsequent dismantling operations can be carried out under proper conditions. All of the combustion products recovered are sent to appropriate reprocessing plants. The method allows high dismantling rates. 1. A method for dismantling solid fuel engines , comprising , for each engine: starting the engine so as to use up all of the propellant that it contains , after which the actual structure of the engine body is taken apart , wherein the engine is started and the propellant is combusted while the engine is mounted on a static test rig and immersed in a tank filled with water.2. The method as claimed in claim 1 , the method including a main sequence having the following steps:{'b': '42', 'a first step which consists in mounting an engine on a static test rig () that holds the engine in a fixed position;'}a second step during which the static test rig, equipped with the engine to be dismantled, is immersed at the bottom of a tank filled with water;a third step of turning on the engine, during which the propellant is used up at the bottom of the tank;a fourth step which consists, after complete combustion of the propellant, in removing the static test rig from the water and dismounting the engine from the test rig;a fifth step which consists in rinsing the body of ...

Preceramic resin formulations, impregnated fibers comprising the preceramic resin formulations, composite materials, and related methods

A preceramic resin formulation comprising a polycarbosilane preceramic polymer, an organically modified silicon dioxide preceramic polymer, and, optionally, at least one filler. The preceramic resin formulation is formulated to exhibit a viscosity of from about 1,000 cP at about 25° C. to about 5,000 cP at a temperature of about 25° C. The at least one filler comprises first particles having an average mean diameter of less than about 1.0 μm and second particles having an average mean diameter of from about 1.5 μm to about 5 μm. Impregnated fibers comprising the preceramic resin formulation are also disclosed, as is a composite material comprising a reaction product of the polycarbosilane preceramic polymer, organically modified silicon dioxide preceramic polymer, and the at least one filler. Methods of forming a ceramic matrix composite are also disclosed.

Preceramic resin formulations, ceramic materials comprising the preceramic resin formulations, and related articles and methods

A preceramic resin formulation comprising a polycarbosilane preceramic polymer and an organically modified silicon dioxide preceramic polymer. A ceramic material comprising a reaction product of the polycarbosilane preceramic polymer and organically modified silicon dioxide preceramic polymer is also described. Articles comprising the ceramic material are also described, as are methods of forming the preceramic resin formulation and the ceramic material.

HYBRID METAL COMPOSITE STRUCTURES, JOINT STRUCTURES, AND RELATED METHODS

A multi-component structure includes a first hybrid metal composite structure, a second hybrid metal composite structure, and a joint structure. The first and second hybrid metal composite structures include layers, each layer comprising a fiber composite material structure including a fiber material dispersed within a matrix material and at least one metal ply located between layers of the layers. The joint structure extends between and connects the first hybrid metal composite structure and the second hybrid metal composite structure. Additionally, the joint structure exerts a clamping force on the first and second hybrid metal composite structures and to reduce gaps between the layers, between the layers and the at least one metal ply, and between the joint structure and the first and second hybrid metal composite structures to less than half a thickness of the at least one metal ply. 1. A multi-component structure , comprising: layers comprising a fiber composite material structure; and', 'at least one metal ply located between the layers; and, 'a first hybrid metal composite structure and a second hybrid metal composite structure, each of the first and second hybrid metal composite structures comprisinga joint structure extending between and connecting the first hybrid metal composite structure and the second hybrid metal composite structure, wherein the joint structure is configured to exert a clamping force on the first and second hybrid metal composite structures to reduce gaps between the layers of the first and second hybrid metal composite structures, between the layers and the at least one metal ply of the first and second hybrid metal composite structures, and between the joint structure and the first and second hybrid metal composite structures to less than half a thickness of the at least one metal ply.2. The multi-component structure of claim 1 , wherein the joint structure comprises:a first connector structure disposed on a first side of the first ...

COMBUSTOR, JET ENGINE, FLYING BODY, AND OPERATION METHOD OF JET ENGINE

A combustor of a jet engine includes a fuel injector, an igniter for igniting a gas mixture of air and fuel, and a flame holder. The igniter is disposed in the flame holder. After an activation of the igniter, the igniter disappears, and a space after the disappearance functions as a flame-holding space. 1. A combustor in which fuel is burned using air taken through an inlet comprising:an injector for injecting the fuel;a flame holder formed in a wall surface of the combustor and configured to keep flame for burning the fuel injected from the injector; andan igniter for igniting a gas mixture of the air and the fuel, wherein the igniter is disposed in the flame holder and configured to disappear to form a flame-holding space in the flame holder.2. The combustor according to claim 1 , wherein the igniter is a rocket motor claim 1 , andwherein the rocket motor is configured to burn and disappear to form the flame-holding space in the flame holder.3. The combustor according to claim 1 , wherein the flame holder formed in the wall surface of the combustor is a frame-holding depression formed in the wall surface of the combustor.4. The combustor according to claim 3 , wherein a maximum length of the frame-holding depression along a direction of a mainstream of the air taken through the inlet is defined as a length L claim 3 , and a maximum depth of the frame-holding depression is defined as a depth D claim 3 , andwherein the length L and the depth D are set such that L>D is satisfied.5. The combustor according to claim 3 , wherein the igniter is installed in a whole of the frame-holding depression.6. The combustor according to claim 3 , wherein the igniter is installed in a part of the frame-holding depression claim 3 , andwherein a part where the igniter is not installed exists in the frame-holding depression.7. The combustor according to claim 1 , wherein at least a part of periphery of the igniter is covered with a barrier member which is distinct from the wall ...

STABLE HYBRID ROCKET TECHNOLOGY

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. 1. A method of producing a stable , efficient combustion in a rocket engine , comprising:generating a core flow of fuel-rich gas generator gases; andinjecting an oxidizer around the core flow of fuel-rich gas generator gases.2. A hybrid rocket engine comprising:a coaxial injector section that is configured to introduce a core flow of fuel-rich gas from a gas-generator;an oxidizer injector section that surrounds the coaxial injector section and that is configured to inject oxidizer around the core flow of fuel-rich gas;a combustion chamber downstream of and in fluid communication with the coaxial injector section and the oxidizer injection section into which a mixture of the fuel-rich gas and oxidizer gas flows and combusts; andan expansion nozzle downstream of and in fluid communication with the combustion chamber through which combustion. products from the combustion chamber are expanded and discharged to produce thrust.3. The hybrid rocket engine of claim 2 , wherein the injector section is fluidly connected to a solid-propellant gas generator that produces the fuel-rich gas claim 2 , and the oxidizer inlet is fluidly connected to an oxidizer tank containing a self-pressurizing oxidizer.4. A hybrid rocket engine claim 2 , comprising:an aerospike nozzle having ...

PROPELLANT GRAIN FOR A SOLID ROCKET MOTOR

The present disclosure relates to propellant grain configuration in solid rocket motors. In one embodiment, the propellant grain is a case-bonded, forward-swept, deep finocyl grain offering significant flexibility in tailoring burn surface area regression profiles to meet different performance requirements even while allowing for high propellant volumetric loading densities. The grain comprises of two or more longitudinal fin cavities with forward swept leading edges, circular-patterned about an axial cavity. 1. A rocket motor , the rocket motor comprising:an internally insulated cylindrical casing with end domes having centrally located apertures at fore-end and aft-end; an axial through-bore running from fore-end to aft-end along the axis of the rocket motor; and', 'a plurality of longitudinal fin cavities circular patterned about the axial through-bore,', 'wherein the plurality of longitudinal fin cavities extend radially outward with forward-swept leading edge and trailing edge., 'a solid propellant grain filled within the casing, comprising2. The rocket motor as claimed in claim 1 , wherein the solid propellant grain further comprises a counter-bore at the aft-end of the rocket motor.3. The rocket motor as claimed in claim 1 , wherein the plurality of longitudinal fin cavities is configured with their leading edges making an acute angle at least one of completely claim 1 , partly and tangentially with the forward motor axis.4. The rocket motor as claimed in claim 1 , wherein the plurality of longitudinal fin cavities disposed about the axial through-bore are dissimilar.5. The rocket motor as claimed in claim 1 , wherein the plurality of longitudinal fin cavities is disposed about the axial through-bore such that the diameter of the largest imaginary circle circumscribing the tips of the fin is greater than the diameter of the aft-end aperture in motor casing and lesser than the inner diameter of the motor casing.6. The rocket motor as claimed in claim 1 , ...

SYSTEMS AND TECHNIQUES FOR LAUNCHING A PAYLOAD

This disclosure describes various techniques and systems for rapid low-cost access to suborbital and orbital space and accommodation of acceleration of sensitive payloads to space. For example, a distributed gas injection system may be used in a ram accelerator to launch multiple payloads through the atmosphere. Additionally or alternatively, multiple projectiles may assemble during flight through the atmosphere to transfer and/or resources to another projectile. 1. A start gun system for accelerating a projectile to ram speed comprising: a start tube configured to selectively inject a pressurant;', 'a control system configured to control staged injection of the pressurant along the start tube; and, 'a distributed injection system comprisinga movable diaphragm configured to increase a relative velocity of the projectile in a medium.2. The start gun system of claim 1 , wherein the movable diaphragm is configured to move towards the projectile through the start tube.3. The start gun system of claim 1 , wherein the selective injection of the pressurant comprises injection into the start gun tube through one or more of gas baffles claim 1 , high speed valves claim 1 , or diaphragms.4. The start gun system of claim 1 , further comprising an onsite distribution system configured to claim 1 , upon request of a user claim 1 , select and load into the start gun system claim 1 , a projectile claim 1 , and launch the selected and loaded projectile.5. The start gun system of claim 1 , further comprising a ram accelerator coupled to the distributed injection system and the moveable diaphragm claim 1 , wherein the ram accelerator is disposed in an underground facility claim 1 , a floating system claim 1 , or a flying vehicle.6. A multiple projectile launch system comprising:a plurality of launch mechanisms configured to accelerate a respective projectile each;a first control system configured to coordinate the acceleration of the respective projectiles in the respective plurality ...

INSULATION, INSULATION PRECURSORS, AND ROCKET MOTORS, AND RELATED METHODS

An insulation material includes a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin. At least one filler, such as hollow glass microspheres and/or carbon fiber is dispersed within the matrix. A rocket motor includes a case, the insulation material within and bonded to the case, and a solid propellant within the case. An insulation precursor includes a silicon carbide precursor resin, a silicon dioxide precursor resin, and the at least one filler. Related methods are also disclosed. 1. An insulation material , comprising:a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin;at least one filler dispersed within the matrix, the at least one filler comprising at least one material selected from the group consisting of a low density filler and an ablation enhancement filler.2. The insulation material of claim 1 , wherein the matrix further comprises a catalyst.3. The insulation material of claim 1 , wherein the insulation material exhibits a thermal conductivity of less than about 0.30 W/mK.4. The insulation material of claim 1 , wherein the insulation material exhibits a density of less than about 0.8 g/cm.5. The insulation material of claim 1 , wherein the insulation material comprises from about 50% to about 85% silicon dioxide by weight.6. The insulation material of claim 1 , wherein the at least one filler comprises hollow glass microspheres.7. The insulation material of claim 6 , wherein the insulation material comprises from about 5% to about 10% hollow glass microspheres by weight.8. The insulation material of claim 6 , wherein the hollow glass microspheres exhibit a mean diameter from about 100 nm to about 5 mm.9. The insulation material of claim 1 , wherein the at least one filler comprises carbon fiber.10. The insulation material of claim 9 , wherein the insulation material comprises from about 1% to about 5% carbon fiber ...

CATALYST CHAMBER WITH A CATALYST BED EMBEDDED THEREIN FOR A MONOPROPELLANT THRUSTER OF A ROCKET ENGINE

A catalyst chamber with a catalyst bed embedded therein for a monopropellant thruster of a rocket engine. The catalyst chamber comprises an inlet having a first cross-sectional area through which a propellant can be introduced into the catalyst chamber and an outlet having a second cross-sectional area through which the propellant and/or resulting reaction products can be introduced into a combustion chamber of the thruster. The outlet is connected to the inlet via a catalyst volume of the catalyst chamber. At least one helical wall member is arranged within the catalyst chamber and is dividing the catalyst volume into two or more segments such that an effective length of the catalyst bed of each segment passed through by the propellant and/or its reaction products is larger than a geometrical length of the catalyst chamber defined between the inlet and the outlet along a direction of extension of the catalyst chamber. 1. A catalyst chamber with a catalyst bed embedded therein for a monopropellant thruster of a rocket engine , comprisingan inlet having a first cross-sectional area through which a propellant can be introduced into the catalyst chamber;an outlet having a second cross-sectional area through which the propellant and/or resulting reaction products can be introduced into a combustion chamber of the thruster wherein the outlet is connected to the inlet via a catalyst volume of the catalyst chamber;at least one helical wall member arranged within the catalyst chamber and dividing the catalyst volume into two or more segments such that an effective length of the catalyst bed of each segment passed through by the propellant and/or its reaction products is larger than a geometrical length of the catalyst chamber which is defined between the inlet and the outlet along a direction of extension of the catalyst chamber.2. The catalyst chamber according to claim 1 , wherein the catalyst chamber has a cylindrical shape having a circular cross-section such that bases ...

Forming machine for forming a hollow body, in particular a casing of a solid propellant engine, and deposit head for such a forming machine

In a machine for forming a hollow body, for example a casing of a solid propellant engine, a tensioned and guided belt comprising a layer of adhesive material and a protective strip is fed from a deposit head, supported by a robot and comprising an unwinding device for unwinding the belt, a winding device for winding the strip, a first pressing roller for pressing the strip and the layer of adhesive material and rotatable about an axis thereof orthogonal to the feed path of the belt and a second pressing roller for pressing the adhesive material and the strip, rotatable about an axis thereof forming an angle other than 90° to the feed path and movable in opposite directions in a direction transversal to feed path. 1. A deposit head for a forming machine for forming hollow bodies; the deposit head comprising a support frame adapted to be coupled to a moving member , motorized supply means and guide means carried by said support frame to feed , along a predefined feeding path , a belt comprising a layer of adhesive material and a protective strip arranged on only one side of said layer of adhesive material , motorized winding means for winding said protective strip and carried by said frame , and a first pressing roller for pushing said strip and said layer of adhesive material toward a deposit surface of said adhesive material and rotatable about an axis thereof that is orthogonal to said feeding path , characterized by further comprising a second pressing roller for pushing said strip and said adhesive material toward said deposit surface; said second pressing roller being carried by said support frame and being rotatable about an axis thereof forming , with said feeding path , an angle other than 90°; first actuation and guide means being interposed between said support frame and said second pressing roller for moving the second pressing roller in opposite directions along a direction transversal to said feeding path.2. A head according to claim 1 , characterized ...

PRECURSOR FORMULATIONS FOR A LINER, A ROCKET MOTOR INCLUDING THE LINER, AND RELATED METHODS

A precursor formulation of a liner comprising a polymer and at least two curatives is disclosed. One of the at least two curatives comprises a curative formulated to preferentially react with the polymer and the other of the at least two curatives comprises a blocked curative formulated to be substantially unreactive with the polymer. A method of lining a rocket motor is also disclosed, as is a rocket motor including the liner. 1. A precursor formulation of a liner , comprising:a polymer; andat least two curatives, one of the at least two curatives comprising a curative formulated to preferentially react with the polymer and the other of the at least two curatives comprising a blocked curative formulated to be substantially unreactive with the polymer.2. The precursor formulation of claim 1 , wherein the polymer comprises hydroxyl functional groups.3. The precursor formulation of claim 1 , wherein the polymer comprises one or more of a polyether claim 1 , a fluorinated polyether claim 1 , a polyurethane claim 1 , an epoxy claim 1 , a polysulfide claim 1 , a polyethylene oxide claim 1 , a polybutadiene claim 1 , or a polyester.4. The precursor formulation of claim 1 , wherein the polymer comprises one or more of a hydroxyl terminated polybutadiene or a hydroxyl terminated polyether.5. The precursor formulation of claim 1 , wherein the polymer comprises from about 10% by weight to about 85% by weight of the precursor formulation.6. The precursor formulation of claim 1 , wherein the curative formulated to preferentially react with the polymer comprises one or more of an aromatic isocyanate or an aliphatic isocyanate.7. The precursor formulation of claim 1 , wherein the curative formulated to preferentially react with the polymer comprises one or more of methylene diphenyl diisocyanate claim 1 , polymeric methylene diisocyanate claim 1 , para-phenylene diisocyanate claim 1 , toluene diisocyanate claim 1 , dimer diisocyanate claim 1 , or isophorone diisocyanate.8. The ...

STRAIN MEASUREMENT DEVICE, A SOLID ROCKET MOTOR INCLUDING SAME, AND RELATED METHODS

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. 1. A strain measurement device , comprising:a reference material;a displacement sensor configured to detect changes in distance between the displacement sensor and the reference material and generate a data signal in response thereto, wherein at least one of the displacement sensor and reference material is coupled with a pre-cured elastomeric material; anda processor operably coupled with the displacement sensor, and configured to determine a strain of another elastomeric material at least partially coupled to the pre-cured elastomeric material based at least in part on the data signal received from the displacement sensor.2. The strain measurement device of claim 1 , wherein the displacement sensor is a Hall-effect sensor claim 1 , and the reference material is a magnet.3. The strain measurement device of claim 1 , wherein the displacement sensor is an Eddy-current sensor claim 1 , and the reference material is selected from the group consisting of a metal and a metal alloy.4. The strain measurement device of claim 1 , wherein the displacement sensor and the reference ...

DEVICE AND METHOD RELATING TO A SENSING DEVICE

Methods, devices, and systems relating to a sensing device are disclosed. A device may comprise a structure including a first surface and a second, opposite surface, wherein the structure comprises one or more segments. Further, the device may include a plurality of sensors disposed on the structure, wherein each segment of the one or more segments comprises a first sensor of the plurality of sensors coupled to the first surface and an associated second sensor of the plurality of sensors coupled to the second surface. Moreover, each sensor of the plurality of sensors may be configured to measure a strain exhibited on an adjacent surface of the structure at an associated segment of the one or more segments. 1. A device , comprising:a structure having a first side and a second side opposite the first side;a first plurality of sensors coupled along the first side of the structure;a second plurality of sensors coupled along the second side of the structure, each sensor of the first plurality of sensors having an associated sensor of the second plurality of sensors, wherein each sensor of the first and second plurality of sensors is configured to measure a strain exhibited by at least a portion of the structure; anda processor configured to determine a characteristic of the structure based on strain data from the first and second plurality of sensors.2. The device of claim 1 , wherein the processor is configured to determine a shape of at least one segment of the structure at a moment in time.3. The device of claim 1 , wherein the processor is configured to determine a shape of the structure at a moment in time by integrating curvature over a length of the structure.4. The device of claim 1 , wherein the processor is configured to determine a shape of the structure as a function of time.5. The device of claim 1 , further comprising an object to which the structure is coupled.6. The device of claim 5 , wherein the object includes:a stationary reference object; anda non- ...

PROPELLANT LOAD, WITH MECHANICALLY REINFORCED LINER/PROPELLANT CONNECTION, AND PREPARATION THEREOF

A propellant load includes a propellant block, containing energetic charges in a crosslinked binder, arranged in a structure having a thermal protection; the crosslinked binder being an energetic binder including a polymer, more polar than hydroxytelechelic polybutadiene (HTPB), which is crosslinked and an energetic plasticizer, the polymer non-crosslinked representing less than 14% of the volume of the propellant block; a bonding layer, based on crosslinked hydroxytelechelic polybutadiene (HTPB), between the thermal protection and the propellant block; and a system for mechanical reinforcement of the bonding layer/propellant block bond, present on at least part of the bonding layer/propellant block interface, including grains embedded in part in the bonding layer and the complementary part thereof being embedded in the propellant block: made of a pyrotechnically inert material, and that has a surface energy greater than 34 mJ/m; and the largest dimension of which is between 0.3 and 5.2 mm. 1. A propellant load comprising:a propellant block, containing energetic charges in a crosslinked binder, arranged in a structure of a substantially cylindrical shape with a sleeve, a rear end and a front end; said structure having a thermal protection attached to its an inner face of the structure, opposite said block;a bonding layer, based on crosslinked hydroxytelechelic polybutadiene (HTPB), between said thermal protection and said propellant block; andmeans for mechanical reinforcement of the a bonding layer/propellant block bond, wherein:said crosslinked binder of said propellant block is an energetic binder comprising a polymer, more polar than hydroxytelechelic polybutadiene (HTPB), which is crosslinked and an energetic plasticizer, said polymer non-crosslinked representing less than 14% of the volume of the propellant block; andsaid means for mechanical reinforcement of the bonding layer/propellant block bond, which are present on at least one portion of the a bonding ...

INSULATION PRECURSORS, ROCKET MOTORS, AND RELATED METHODS

An insulation material includes a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin. At least one filler, such as hollow glass microspheres and/or carbon fiber is dispersed within the matrix. A rocket motor includes a case, the insulation material within and bonded to the case, and a solid propellant within the case. An insulation precursor includes a silicon carbide precursor resin, a silicon dioxide precursor resin, and the at least one filler. Related methods are also disclosed. 1. A rocket motor , comprising:a case; a matrix comprising a reaction product formed from a silicon carbide precursor resin and a silicon dioxide precursor resin; and', 'at least one filler dispersed within the matrix, the at least one filler comprising at least one material selected from the group consisting of a low density filler and an ablation enhancement filler; and, 'an insulation material within the case, the insulation material comprisinga solid propellant within the case.2. The rocket motor of claim 1 , further comprising a nozzle secured to the case.3. The rocket motor of claim 1 , wherein the insulation material is bonded to the case with a silicone adhesive.4. An insulation precursor claim 1 , comprising:a silicon carbide precursor resin;a silicon dioxide precursor resin; andat least one filler material selected from the group consisting of a low density filler and an ablation enhancement filler.5. The insulation precursor of claim 4 , wherein the silicon dioxide precursor resin comprises an organically modified silicon dioxide preceramic polymer.6. The insulation precursor of claim 4 , further comprising a catalyst.7. The insulation precursor of claim 4 , further comprising an adhesion promoter.8. The insulation precursor of claim 4 , wherein the silicon dioxide precursor resin exhibits a density from about 0.95 g/cmto about 1.05 g/cm.9. The insulation precursor of claim 4 , wherein the silicon carbide ...

EFFECTOR HEALTH MONITOR SYSTEM AND METHODS FOR SAME

An effector health monitor system is configured for coupling with an energetic component. The effector health monitor system includes a characteristic sensor suite including at least first and second characteristic sensors. The first characteristic sensor is proximate to the energetic component and configured to measure a failure characteristic of the energetic component. The second characteristic sensor is configured to measure at least one environmental characteristic proximate to the energetic component. A communication hub is coupled with the first and second characteristic sensors, and is configured to communicate the measured failure and environmental characteristics outside of an effector body. A failure identification module compares the measured failure characteristic with a failure threshold and identifies a failure event. A failure model generation module logs the at least one measured environmental characteristic preceding the identified failure event with the identified failure event and generates a failure model including updating the failure model. 1. An effector comprising:an effector body including a rocket motor having a solid propellant grain; [ at least the first characteristic sensor is engaged with the solid propellant grain and configured to measure a failure characteristic of the solid propellant grain; and', 'the second characteristic sensor is configured to measure at least one environmental characteristic proximate to the solid propellant grain;, 'a characteristic sensor suite including at least first and second characteristic sensors coupled with the effector, 'a communication hub coupled with at least the first and second characteristic sensors, the communication hub is configured to communicate the measured failure and environmental characteristics outside of the effector body;', 'a failure identification module configured to compare at least the measured failure characteristic with a failure threshold and identify a failure event based ...

VARIABLE BURN-RATE SOLID ROCKET MOTOR IGNITION METHOD

A solid rocket motor uses at least one thermally conductive wire or at least one pair of electrically conductive wires to increase a burn surface area of a propellant grain and thus a thrust of the rocket motor. The rocket motor includes a pulse chamber containing a burnable propellant grain, a propellant inhibited center bore bonded to surfaces of the burnable propellant grain, and at least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore. The conductive wire is configured for passive or active activation to ignite the propellant inhibited center bore that subsequently burns in the variable regions. The thermally conductive wire is formed of a refractory metal or refractory alloy material that enables the entire length of the wire to be heated simultaneously or nearly simultaneously when the wire is passively activated. 1. A rocket motor comprising:a pulse chamber containing a burnable propellant grain;a propellant inhibited center bore bonded to surfaces of the burnable propellant grain; andat least one conductive wire coupled to the burnable propellant grain and arranged in variable regions along the propellant inhibited center bore, the at least one conductive wire being configured for passive or active activation to ignite and ablate the propellant inhibited center bore that subsequently burns in the variable regions, wherein the at least one conductive wire is interposed between the propellant inhibited center bore and the propellant grain to increase a burn surface area of the burnable propellant grain via ablation of the propellant inhibited center bore.2. The rocket motor according to claim 1 , wherein the at least one conductive wire is thermally conductive and passively activated by a burn front of the burnable propellant grain that directly impinges the at least one conductive wire.3. The rocket motor according to claim 2 , wherein an entire length of the at least one ...

ELECTRICAL DEVICE WITH SHUNT, AND RECEPTACLE

An electrical device has device electrical contacts that are initially shunted together, to prevent accidental triggering or damage to the device, such as by electrostatic forces. The device is configured to be inserted into a receptacle, with parts of the receptacle disengaging the shunt and making electrical connection within the receptacle, such as with a shunt cutter. The receptacle may also include a pair of receptacle electrical contacts the electrically connect to the device electrical contacts. The configuration, where the shunt is only cut as part of the installation process, enables safer handling of initially-shunted devices, and can also facilitate making blind electrical connections. Making blind connection directly with parts of the receptacle also avoids the need to thread wires through the electrical receptacle and make electrical connections in another way. 1. An installation , comprising:an electrical device; anda receptacle that receives and electrically connects with the electrical device; a pair of device electrical contacts; and', 'a shunt electrically connecting the pair of device electrical contacts;, 'wherein the electrical device includes a pair of receptacle electrical contacts; and', 'a cutter; and, 'wherein the receptacle includeswherein, when the electrical device is inserted into the receptacle, the cutter severs the shunt, breaking the electrical connection between the device electrical contacts, and allowing electrical connection between the device electrical contacts and the receptacle electrical contacts.2. The installation of claim 1 , wherein the cutter is a protrusion from the receptacle claim 1 , between the receptacle electrical contacts.3. The installation of claim 2 , wherein the cutter protrudes further than the receptacle electrical contacts from a surface of the receptacle.4. The installation of claim 1 , wherein the electrical device is an energetic device that includes an energetic material that is electrically actuated ...

PRECURSOR COMPOSITIONS FOR AN INSULATION AND INSULATED ROCKET MOTORS

A precursor composition comprising, before curing, ethylene propylene diene monomer (EPDM), zinc oxide, silica, polymerized 1,2-dihydro-2,2,4-trimethylquinoline, a solid chlorinated paraffin, stearic acid, a five carbon petroleum hydrocarbon, trimethylolpropane trimethacrylate, and a peroxide. A rocket motor including a reaction product of the precursor composition and a method of insulating a rocket motor. 1. A precursor composition , comprising , before curing:ethylene propylene diene monomer (EPDM), zinc oxide, silica, a solid chlorinated paraffin, and a curative.2. The precursor composition of claim 1 , wherein the precursor composition comprises from about 4.2 parts to about 9 parts of the solid chlorinated paraffin.3. The precursor composition of claim 1 , wherein the precursor composition comprises from about 0.35 part to about 0.75 part of the stearic acid.4. The precursor composition of claim 1 , wherein the precursor composition comprises from about 2.1 parts to about 4.5 parts of the zinc oxide.5. The precursor composition of claim 1 , wherein the precursor composition comprises from about 21 parts to about 45 parts of the silica.6. The precursor composition of claim 1 , further comprising trimethylolpropane trimethacrylate.7. The precursor composition of claim 1 , wherein the curative comprises a peroxide.8. The precursor composition of claim 1 , wherein the precursor composition comprises less than about 12 parts of the curative.9. A precursor composition claim 1 , comprising claim 1 , before curing:ethylene propylene diene monomer (EPDM), zinc oxide, silica, a solid chlorinated paraffin, stearic acid, an antioxidant, a plasticizer, a coagent, and a peroxide.10. The precursor composition of claim 9 , wherein the EPDM comprises a non-conjugated diene and the EPDM comprises a diene content of from about 1% by weight (wt %) to about 10 wt %.11. The precursor composition of claim 9 , wherein the EPDM comprises an ethylene content of between about 40 wt % ...

Brazing method

According to this brazing method, a first base member having a non-plated surface, a metal layer for functioning as a diffusion barrier layer, a brazing foil, and a second base member having a surface are arranged in this order so that the non-plated surface of the first base member and the surface of the second base member are faced with each other. The first base member and the second base member are brazed by using the brazing foil. The cost of providing a diffusion barrier layer between the first base member and the brazing foil is thereby reduced.

THERMALLY INTITIATED VARIABLE VENTING SYSTEM

A thermally initiated variable venting system may comprise a first linear shape charge (LSC) coupled to a first sensor and a second LSC coupled to a second sensor. An upper apex of the second LSC may be disposed within a lower apex of the first LSC. The output of the system may vary depending on whether the event is fast cook-off (FCO) or slow cook-off (SCO). 1. A thermally initiated variable venting system , comprising:a first linear shape charge (LSC) coupled to a first sensor; anda second LSC coupled to a second sensor;wherein the first LSC overlaps the second LSC;the first sensor is configured to activate to ignite the first LSC in response to at least a portion of the first sensor reaching a first temperature;the second sensor is configured to activate to ignite the second LSC in response to at least a portion of the second sensor reaching a second temperature;the thermally initiated variable venting system configured to project a molten jet in response to at least one of the first LSC being ignited by the first sensor and the second LSC being ignited by the second sensor.2. The thermally initiated variable venting system of claim 1 , wherein the molten jet is greater in response to the first LSC being ignited by the first sensor.3. The thermally initiated variable venting system of claim 2 , wherein the molten jet is configured to cut a slot through a motor case in response to the first LSC being ignited by the first sensor.4. The thermally initiated variable venting system of claim 2 , wherein the molten jet is configured to cut a trench into a motor case in response to the second LSC being ignited by the second sensor.5. The thermally initiated variable venting system of claim 2 , wherein at least one of a velocity claim 2 , a pressure claim 2 , and a temperature of the molten jet is greater in response to the first LSC being ignited by the first sensor than in response to the second LSC being ignited by the second sensor.6. The thermally initiated variable ...

Rocket motor with concentric propellant structures for shock mitigation

A solid rocket motor includes a first solid propellant and a second solid propellant at least partially surrounding the first solid propellant. The second solid propellant is resistant to fragment impact and the first solid propellant has a higher impulse than the second solid propellant.

Semipreg with thermoplastic toughened novolac-based epoxy resin matrix

A semipreg that can be cured/molded to form aerospace composite parts including rocket booster casings. The semipreg includes a fibrous layer and a resin layer located on one side of the fibrous layer. The resin layer includes an epoxy component that is a combination of a hydrocarbon epoxy novolac resin and a trifunctional epoxy resin and optionally a tetrafunctional epoxy resin. The resin matrix includes polyethersulfone as a toughening agent and a thermoplastic particle component.

MULTI-PULSE SOLID ROCKET MOTOR IGNITION METHOD

A rocket motor has an electrically operated propellant initiator for a propellant grain that includes an electrode arrangement configured to concentrate an electric field at an ignition electrode for igniting an electrically operated propellant. The rocket motor includes a combustion chamber containing at least one propellant grain and an electrically operated propellant initiator operatively coupled to the propellant grain to initiate combustion of the propellant grain. The electrically operated propellant initiator includes the electrically operated propellant and at least one pair of electrodes configured to ignite the electrically operated propellant. The pair of electrodes includes a ground plane electrode and an ignition electrode. When an electrical input is applied to the electrically operated propellant initiator, the electric field is concentrated at the ignition electrode to ignite the electrically operated propellant at the location where the ignition electrode is arranged. 1. A rocket motor comprising:a combustion chamber containing at least one propellant grain; andan electrically operated propellant initiator operatively coupled to the at least one propellant grain to initiate combustion of the at least one propellant grain, the electrically operated propellant initiator including an electrically operated propellant and at least one pair of electrodes arranged to ignite the electrically operated propellant, the at least one pair of electrodes including a ground plane electrode and an ignition electrode at which an electric field is concentrated to ignite the electrically operated propellant,wherein the ground plane electrode extends along a first surface area of the at least one propellant grain that is larger than a second surface area of the at least one propellant grain along which the ignition electrode extends, wherein a ratio of the first surface area to the second surface area is at least 2:1.2. The rocket motor according to claim 1 , wherein ...

MANDREL ASSEMBLY AND METHOD OF MANUFACTURING SOLID ROCKET PROPELLANT GRAIN USING THE SAME

The present disclosure relates to a dismantleable mandrel assembly and a method of molding solid propellant grains with deep fin cavities whose major transverse dimensions are larger than casing opening dimensions in a monolithic rocket motor. The mandrel assembly comprises a base mandrel, a core mandrel insertable into the base mandrel and a plurality of fin molds attachable onto the base mandrel in a circular pattern about the motor axis. The plurality of longitudinal fin cavities is configured with forward swept leading and trailing edges. The manufacturing technique involves assembling and disassembling the mandrel components before propellant casting and after propellant curing respectively in a specific sequence. With minimum number of components and critical joints the method assures reduced quantum of explosive hazard in propellant grain manufacturing for high performance solid rocket motors. 1. A mandrel assembly for manufacturing a solid propellant grain of a rocket motor , said mandrel assembly comprising:a base mandrel removably connectable to aft-end opening of rocket motor casing capable of forming an aft-end counter bore in the propellant grain of the rocket motor;a core mandrel removably connectable to the base mandrel, wherein the core mandrel is capable of forming a longitudinal axial cavity in the propellant grain; anda plurality of fin molds, removably connectable to the base mandrel, wherein the plurality of fin molds comprises a forward-swept leading edge and a forward-swept trailing edge to form a plurality of forward-swept longitudinal hollow fins circular-patterned about the axial cavity in the solid propellant grain of the rocket motor.2. The mandrel assembly as claimed in claim 1 , wherein the base mandrel is a hollow axisymmetric structure configured with a port portion comprising a plurality of guides disposed in a circular pattern on outer circumference of the base mandrel.3. The mandrel assembly as claimed in claim 1 , wherein the ...

INTEGRATED THRUSTER

A thruster has an additively-manufactured housing that includes an integrally-formed nozzle with a burst disk in it. The housing is part of a casing that surrounds and encloses a propellant that is burned to produce pressurized gases that burst the burst disk and produce thrust. The thruster may be placed in a receptacle that defines a recess for receiving the thruster. The receptacle also may be additively manufactured. The thruster and the recess both may be cylindrical, with the housing being closely fit with the cylindrical walls of the receptacle. This may allow some of the structural loads on the housing, such as loads produced by the combustion of the propellant, to be transferred to the adjoining walls of the receptacle. This enables the housing to have less structural strength than if it were to have to contain the pressure from the propellant all on its own. 1. A thruster comprises:a propellant; anda casing enclosing the propellant;wherein the casing is an additively-manufactured housing; andwherein the housing defines an integral nozzle through which pressurized gases exit the casing when the propellant is burned.2. The thruster of claim 1 ,wherein the housing also includes a burst disk in the nozzle;wherein the burst disk is additively-manufactured as an integral part of the additively-manufactured housing.3. The thruster of claim 1 ,wherein the housing includes a flat end portion and an annular portion that are additively-manufactured together as a single piece;wherein the flat end portion closes off an end of the annular portion; andwherein the nozzle is in the flat end portion.4. The thruster of claim 3 , wherein the housing includes a closure that closes off an end of the annular portion that is opposite the end of the annular portion that is closed off by the flat end portion.5. The thruster of claim 4 , wherein the closure is welded to the annular portion.6. The thruster of claim 1 , wherein the housing includes a closure that forms a welded seal ...

Solid rocket motor with vortex inducing feature

A solid rocket motor includes a propellant grain structure defining an axial bore and a vortex inducing feature.

SOLID ROCKET MOTORS INCLUDING FLIGHT TERMINATION SYSTEMS, AND RELATED MULTI-STAGE SOLID ROCKET MOTOR ASSEMBLIES AND METHODS

A solid rocket motor comprises a pressure vessel, a solid propellant structure within the pressure vessel, and a flight termination system overlying the pressure vessel. The flight termination system comprises a shaped charge configured and positioned to effectuate ignition of an inner portion of the solid propellant structure and a reduction in an ability of the pressure vessel to withstand a change in internal pressure. Another solid rocket motor, a multi-stage rocket motor assembly, and a method of destroying a launch vehicle in flight are also described. 1. A solid rocket motor , comprising:a pressure vessel;a solid propellant structure within the pressure vessel; anda flight termination system overlying the pressure vessel and comprising a shaped charge configured and positioned to effectuate ignition of an inner portion of the solid propellant structure and a reduction in an ability of the pressure vessel to withstand a change in internal pressure.2. The solid rocket motor of claim 1 , further comprising:a bore longitudinally extending through the solid propellant structure; anda slot in communication with an end of the bore and exhibiting a radial end proximate a sidewall of the pressure vessel, the shaped charge configured and positioned to effectuate ignition of the solid propellant structure at the end of the bore.3. The solid rocket motor of claim 2 , wherein the shaped charge is configured and positioned to cut through portions of the pressure vessel and the solid propellant structure overlying the radial end of the slot and to at least partially cut through additional portions of the pressure vessel not overlying the radial end of the slot.4. The solid rocket motor of claim 2 , wherein a portion of the shaped charge is positioned over and laterally aligned with the radial end of the slot.5. The solid rocket motor of claim 1 , wherein the shaped charge comprises a linear shaped charge.6. The solid rocket motor of claim 1 , wherein the flight termination ...

RING-SHAPED BOOSTER ROCKET

A rocket booster has an annular shape, with a casing defining an annular space therewithin, and a solid rocket fuel in the annular spacing. The casing may itself at least in part define an annular gap that functions as a nozzle for the rocket booster, with protruding tabs on the casing aiding in maintaining a uniform height of the annular gap. The rocket booster may be mechanically coupled to an object protruding from the back of a fuselage of a flight vehicle, such as a missile. For example, the rocket booster may be placed around an aft turbojet nozzle of the flight vehicle. This allows the rocket booster to be used in situations where primary propulsion must be running both before and after (and perhaps during) the firing of the rocket booster. 1. (canceled)2. The booster rocket of claim 6 , wherein the casing is made of metal.3. The booster rocket of claim 6 , wherein the casing is a single unitary part.4. The booster rocket of claim 6 , wherein the casing is multiple casing parts.5. The booster rocket of claim 4 ,wherein the multiple casing parts include an inner casing part and an outer casing part; andwherein the inner casing part and the outer casing part are threadedly coupled together.6. A booster rocket comprising:an annular casing defining an annular space therewithin, and having a central opening; anda solid rocket fuel in the annular space;wherein the annular casing defines an annular gap that acts as a nozzle for the booster rocket; andwherein the booster rocket is a capable of being placed around and installed around a separate object.7. The booster rocket of claim 6 , wherein the casing includes protruding tabs that maintain the annular gap.8. The booster rocket of claim 6 , wherein the casing includes an inner part and an outer part claim 6 , with the outer part including a cylindrical forward section claim 6 , and an inwardly-sloped aft section.9. The booster rocket of claim 8 , wherein the inner part of the casing includes a cylindrical forward ...

RING-SHAPED BOOSTER ROCKET

A rocket booster has an annular shape, with a casing defining an annular space therewithin, and a solid rocket fuel in the annular spacing. The rocket booster also includes one or more nozzle pieces, mechanically coupled to the casing, that define one or more nozzles at the aft side of the rocket booster. The rocket booster may be mechanically coupled to an object protruding from the back of a fuselage of a flight vehicle, such as a missile. For example, the rocket booster may be placed around an aft turbojet nozzle of the flight vehicle. This allows the rocket booster to be used in situations where primary propulsion must be running both before and after (and perhaps during) the firing of the rocket booster. 1. (canceled)2. A booster rocket comprising:an annular casing defining an annular space therewithin, and having a central opening;a solid rocket fuel in the annular space; andone or more nozzle pieces mechanically coupled to the annular casing, defining one or more nozzles at an aft end of the annular casing;wherein the one or more nozzle pieces includes an annular nozzle piece that defines an annular nozzle; andwherein the booster rocket is a capable of being placed around and installed around a separate object.3. The booster rocket of claim 2 , further comprising a seal between the annular nozzle piece and the annular casing.4. The booster rocket of claim 2 , wherein the annular nozzle piece defines the annular nozzle in combination with an inner nozzle insert that is attached to the annular casing.5. The booster rocket of claim 4 , wherein the annular nozzle piece and the inner nozzle insert together constitute a throat insert set.6. The booster rocket of claim 2 , wherein the annular nozzle piece has protrusions protruding inward from an inner edge claim 2 , the protrusions facilitating maintaining an annular gap of the annular nozzle.7. The booster rocket of claim 2 , wherein the annular nozzle piece is made of a metallic material.8. The booster rocket of ...

Precursor compositions for an insulation, insulated rocket motors, and related methods

A precursor composition comprising, before curing, ethylene propylene diene monomer (EPDM), an aramid, and a carbon material comprising carbon nanotubes, graphite, or a combination thereof. A rocket motor including a reaction product of the precursor composition and a method of insulating a rocket motor are also disclosed.

IGNITION SAFETY DEVICE FOR ROCKET MOTOR

Provided is an ignition safety device for a rocket motor. The device includes an ignition circuit part including a main body and a control unit configured to generate an initiation signal on the basis of a specific signal, and an initiation part mounted at one end of the ignition circuit part and including at least one high voltage initiator electrically connected to the ignition circuit part. The initiation part includes a housing having at least one reception space for receiving the at least one high voltage initiator, and the housing and the main body are coupled by welding. 16-. (canceled)7. An ignition safety device for a rocket motor , the ignition safety device comprising:an ignition circuit part comprising a main body and a control unit configured to generate an initiation signal on the basis of a specific signal; andan initiation part mounted at one end of the ignition circuit part and comprising at least one high voltage initiator electrically connected to the ignition circuit part,wherein the initiation part comprises a housing having at least one reception space for receiving the at least one high voltage initiator,wherein the housing and the main body are coupled by welding,wherein the high voltage initiator is received at one end of the housing coupled to the main body, andwherein the housing further comprises at least one inner hole corresponding to the high voltage initiator.8. The ignition safety device of claim 7 , further comprising an O-ring disposed at an inner surface of the housing constituting the reception space and an inner surface of the high voltage initiator to prevent fluid movement claim 7 ,wherein a recess area for inserting the O-ring is disposed at an outer surface of the high voltage initiator.9. The ignition safety device of claim 7 , wherein when a plurality of high voltage initiators are received in the housing claim 7 , the plurality of reception spaces are disposed apart from each other to separately store the plurality of ...

COMMON BULKHEAD FOR A PRESSURE VESSEL

The invention lies in the field the management of pressures and relates to a common bulkhead for a pressure vessel having two chambers, the common bulkhead being intended to be positioned between a first chamber and a second chamber of the pressure vessel and configured to withstand a first predetermined pressure in the first chamber and to allow a fluid from the second chamber to flow above a second predetermined pressure, wherein it comprises: a metallic basic structure comprising a first face intended to be positioned facing towards the first chamber, a second face intended to be positioned facing towards the second chamber, a plurality of through-openings between the first face and the second face having a polygonal-type pattern in section, an external frame at its periphery, a first metallic cap superposed on the first face covering the plurality of through-openings. 1. A Common bulkhead for a pressure vessel having two chambers , the common bulkhead being intended to be positioned between a first chamber and a second chamber of the pressure vessel and configured to withstand a first predetermined pressure in the first chamber and to allow a fluid from the second chamber to flow above a second predetermined pressure , comprising:a basic structure comprising a first face intended to be positioned facing towards the first chamber, a second face intended to be positioned facing towards the second chamber, a plurality of through-openings between the first face and the second face,a first metallic cap superposed on the first face covering the plurality of through-openings,wherein the plurality of through-openings have a polygonal section.2. The common bulkhead according to claim 1 , wherein a polygonal section has an edge substantially parallel to an edge of a polygonal section adjacent thereto.3. The common bulkhead according to claim 1 , wherein at least one of the plurality of through-openings comprises a means for reducing the stress concentrations at at least ...

HYBRID METAL COMPOSITE STRUCTURES, ROCKET CASES, AND RELATED METHODS

A method of forming a hybrid metal composite structure including at least one metal ply. The method includes forming at least one metal ply, forming the at least one metal ply comprising forming at least one perforation in the at least one metal ply, abrasively blasting at least one surface of the at least one metal ply to coarsen the at least one surface of the metal ply, and exposing the at least one metal ply to at least one of an acid or a base. The method further includes disposing at least one fiber composite material structure adjacent the at least one metal ply. Related methods of forming a portion of a rocket case and related hybrid metal composite structures are also disclosed. 1. A method of forming a hybrid metal composite structure , the method comprising: forming at least one perforation in the at least one metal ply;', 'abrasively blasting at least one surface of the at least one metal ply to coarsen the at least one surface of the at least one metal ply; and', 'exposing the at least one metal ply to at least one of an acid or a base; and, 'forming at least one metal ply, forming the at least one metal ply comprisingdisposing at least one fiber composite material structure adjacent the at least one metal ply.2. The method of claim 1 , further comprising selecting the fiber composite material structure to comprise a carbon fiber composite material.3. The method of claim 1 , wherein abrasively blasting at least one surface of the at least one metal ply comprises forming a surface of the metal ply having a surface roughness from about 0.5 μm Rto about 4.0 μm R.4. The method of claim 1 , wherein exposing the at least one metal ply to at least one of an acid or a base comprises exposing the at least one metal ply to one of sulfuric acid claim 1 , nitric acid claim 1 , hydrochloric acid claim 1 , hydrofluoric acid claim 1 , or combinations thereof.5. The method of claim 1 , further comprising selecting the at least one metal ply to comprise titanium.6. The ...

ELECTRICALLY OPERATED PROPELLANT FOR SOLID ROCKET MOTOR THRUST MANAGEMENT

Electrically operated propellant is used to supplement the thrust provided by solid rocket motor (SRM) propellant to manage thrust produced by a SRM. The gas produced by burning the electrically operated propellant may be injected upstream of the nozzle to add mass and increase chamber pressure Pc, injected at the throat of the nozzle to reduce the effect throat area At to increase chamber pressure Pc or injected downstream of the throat to provide thrust vector control or a combination thereof. Certain types of electrically operated propellants can be turned on and off provided the chamber pressure Pc does not exceed a self-sustaining threshold pressure eliminating the requirement for physical control valves. 1. A solid rocket motor , comprising:a combustion chamber;a nozzle coupled to the combustion chamber, said nozzle having a throat with an effective throat area At;one or more segments of solid rocket motor (SRM) propellant within the combustion chamber;a thermal ignition coupled to at least one of the one or more segments of SRM propellant, each said segment of SRM propellant, once ignited, burns to completion to produce pressurized gas in the combustion chamber at a chamber pressure Pc that flows through the throat of the nozzle to produce thrust;an electrically operated propellant; andan electrical ignition configured to apply an electrical input to the electrically operated propellant, wherein in an ignition condition the electrically operated propellant burns to produce gas to (a) add mass into the chamber upstream of the nozzle throat to increase the chamber pressure Pc, (b) reduce the effective throat area At to increase chamber pressure Pc or (c) to inject mass downstream of the nozzle throat to provide thrust vector control (TVC) or a combination thereof.2. The solid rocket motor of claim 1 , wherein said electrically operated propellant has a self-sustaining threshold pressure at which the propellant once ignited cannot be extinguished and below which ...

SOLID ROCKET MOTOR WITH BARRIER

A solid rocket motor includes a propellant grain and a barrier shielding at least a portion of the grain. The barrier is impermeable to water, oxygen, nitrogen, and volatile solid propellant species. 1. A solid rocket motor comprising:a propellant grain; anda barrier surrounding at least a portion of the grain, wherein the barrier is impermeable to water, oxygen, nitrogen and volatile solid propellant species.2. The solid rocket motor as recited in claim 1 , wherein the barrier includes an aluminized polymer material.3. The solid rocket motor as recited in claim 1 , wherein the barrier is a hermetic barrier.4. The solid rocket motor as recited in claim 1 , wherein the barrier includes a barrier layer lining the bore.5. The solid rocket motor as recited in claim 4 , wherein the barrier layer includes a material selected from the group consisting of polymeric material claim 4 , ceramic material claim 4 , metallic material claim 4 , and combinations thereof.6. The solid rocket motor as recited in claim 4 , wherein the barrier layer includes a substrate layer and a metallic layer disposed on the substrate layer.7. The solid rocket motor as recited in claim 6 , wherein the bore defines an axis claim 6 , and the substrate layer is radially outwards of the metallic layer.8. The solid rocket motor as recited in claim 1 , further comprising an ignition system that includes an ignition cord between the barrier and the propellant grain.9. The solid rocket motor as recited in claim 1 , further comprising an ignition system operable to ignite the propellant grain claim 1 , the ignition system including a multi-metallic ignition body having at least two metallic elements in contact with each other and a fluorine-containing body in contact with the multi-metallic ignition body.10. The solid rocket motor as recited in claim 9 , wherein the at least two metallic elements include aluminum and palladium.11. The solid rocket motor as recited in claim 9 , wherein the barrier includes ...

Rocket delay apparatuses, systems and methods

Delay tools, systems and methods for achieving a selection of alternative delay times, a tool of which including a body, a drill bit operable relative to the body and a knob operably connected to the drill bit, and operably disposed relative to the body for engagement of the body with a rocket motor bulkhead and the drill both relative to a delay to provide for achieving a selection of alternative delay times.

TEST METHOD FOR TESTING A SOLID-PROPELLANT ROCKET ENGINE, SOLID-PROPELLANT ROCKET ENGINE AND SYSTEM FOR IMPLEMENTING THE METHOD

A solid-propellant rocket engine () has a casing () and a thermal protection () internally coating the casing and delimiting a housing (), which contains a mass of solid propellant (); the thermal protection has a fixed portion () and at least one movable portion () that adheres to the mass of solid propellant () and can be moved from a back position to a forward position with respect to the fixed portion () through a thrust system obtained by pressuring a chamber provided by installing a membrane between the fixed portion and the movable portion ; the engine is tested by verifying the adhesion of the mass of solid propellant () to the movable portion () after having moved the movable portion () to the forward position by means of a thrust directed from the fixed portion towards the mass of solid propellant (). 1. A test method for testing a solid-propellant rocket engine comprising a mass of solid propellant , a casing extending along an axis and said casing comprising two annular end portions opposite to each other along said axis and a heat protection arranged to coat an inner surface of said casing and said heat protection delimiting a housing containing said mass of solid propellant , said heat protection including a coating portion comprising a fixed portion and at least one movable portion , said fixed portion coating at least one of said two annular end portions , said at least one movable portion having a face that adheres to said mass of solid propellant and said at least one movable portion is movable between a back position and a forward position , said at least one movable portion being closer to said fixed portion in said back position and said at least one movable portion being more distant from said fixed portion in said forward position , the method comprising the steps of:moving said at least one movable portion to said forward position by exerting, on said at least one movable portion, a thrust oriented from said fixed portion towards said mass of ...

Device for connecting two segments of a propelling nozzle

The invention relates to the field of propulsion nozzles, and in particular to a device ( 105 ) for connecting together first and second segments ( 103 a, 103 b ) of a propulsion nozzle that are made of thermally dissimilar materials. The device ( 105 ) comprises at least one pin ( 106 ) and an eccentric bushing ( 107 ). The pin ( 106 ) presents both a first axisymmetric surface ( 106 a ) that is to be housed in a radial orifice ( 108 ) of the first nozzle segment ( 103 a ) and also a second axisymmetric surface ( 106 b ) that is eccentric relative to said first axisymmetric surface ( 106 a ).

MANUFACTURING PROCESS FOR MAKING A DOME ELEMENT PROVIDED WITH THERMAL PROTECTION FOR A SOLID PROPELLANT ROCKET ENGINE

For producing a dome-shaped element () provided with thermal protection for a solid propellant rocket engine, a coupling annular body () is arranged in a mold () and has a surface () that is clean and activated, by an atmospheric-pressure plasma treatment, before depositing a primer layer () and an adhesive layer () on the surface (); ablative material is then automatically applied to the adhesive layer and to an area () of the mold () so as to form a series of superimposed layers (). 2. The method of claim 1 , wherein the plasma treatment is carried out after having arranged said coupling annular body in said mold.3. The method of claim 2 , wherein the plasma treatment is carried out by means of a movable head of a robot.4. The method of claim 1 , wherein the ablative material is applied in form of a web by means of a movable head of a robot.5. The method of claim 4 , wherein said web is stored in form of one or more reels claim 4 , carried by said robot.6. The method of claim 1 , comprising the step of heating the ablative material before and/or during the application onto said fixing intermediate layer.7. The method of claim 1 , wherein said fixing intermediate layer is deposited by spraying.8. The method of claim 7 , wherein said fixing intermediate layer is deposited by means of a movable head of a robot and is defined by at least one raw material contained in a tank carried by said robot.9. A station for producing a dome-shaped element provided with thermal protection for a solid propellant rocket engine; the station comprising:a mold;a robot comprising a movable head;a device for atmospheric-pressure plasma treatment, carried by said movable head.10. The station of claim 9 , further comprising an application device carried by said movable head for applying the ablative material in form of web.11. The station of claim 10 , wherein said application device comprises a heating system for heating said web.12. The station of claim 10 , wherein said application ...

ROCKET MOTOR AUXILIARY POWER GENERATION UNIT SYSTEMS AND METHODS

A method for generating electric power for a rocket system includes burning a primary solid propellant grain to create a primary high pressure gas for providing thrust to the rocket, opening a first valve to divert a portion of the high pressure gas to an auxiliary solid propellant grain for igniting the auxiliary solid propellant grain, wherein the auxiliary solid propellant grain is disposed in a housing separate from the primary solid propellant grain, and burning the auxiliary solid propellant grain to create an auxiliary high pressure gas for turning a turbine. The method further includes driving a generator with the turbine and generating an electric power with the generator. 1. A system connectable with a rocket , comprising:a primary motor comprising a primary solid propellant grain configured to burn to create a primary high pressure gas;an auxiliary gas generator comprising an auxiliary solid propellant grain disposed in a housing separate from the primary solid propellant grain;a first valve;an electric generator; anda turbine coupled to the electric generator; in response to the first valve moving to an open position, the primary motor is in fluid communication with the auxiliary gas generator for igniting the auxiliary solid propellant grain; and', 'in response to the auxiliary solid propellant grain being ignited by the primary high pressure gas, the auxiliary solid propellant grain is configured to burn to create an auxiliary high pressure gas., 'wherein,'}2. The system of claim 1 , further comprising a second valve for metering the auxiliary high pressure gas to the turbine.3. The system of claim 2 , wherein the auxiliary high pressure gas is configured to be directed to the turbine in response to the second valve moving to an open position.4. The system of claim 3 , further comprising a third valve for dumping pressure from the housing to extinguish the auxiliary solid propellant grain.5. The system of claim 1 , wherein the auxiliary solid ...

PROPELLANT

A propellant may comprise a primary oxidizer having a density of greater than or equal to 2.7 grams per cubic centimeter; a secondary oxidizer having a density of less than 2.7 grams per cubic centimeter; and/or a fuel comprising at least one of zirconium metal, tin metal, titanium metal, aluminum metal, magnesium metal, or a metal hydride. The propellant may be substantially free of lead. 1. A propellant , comprising:a primary oxidizer having a density of greater than or equal to 2.7 grams per cubic centimeter; anda fuel comprising at least one of zirconium metal, tin metal, tungsten metal, zinc metal, titanium metal, aluminum metal, magnesium metal, magnalium metal alloy, or a metal hydride,wherein the propellant is substantially free of lead.2. The propellant of claim 1 , wherein the primary oxidizer comprises at least one of potassium iodate claim 1 , potassium periodate claim 1 , cesium iodate claim 1 , cesium periodate claim 1 , cupric iodate claim 1 , copper periodate claim 1 , or ammonium iodate.3. The propellant of claim 2 , comprising a secondary oxidizer having a density of less than 2.7 grams per cubic centimeter claim 2 , wherein the secondary oxidizer comprises at least one of potassium perchlorate or ammonium perchlorate.4. The propellant of claim 3 , comprising about 50% to about 90% by weight the primary oxidizer.5. The propellant of claim 4 , comprising about 1% to about 50% by weight the secondary oxidizer.6. The propellant of claim 5 , comprising 2% to 30% by weight the fuel.7. The propellant of claim 1 , further comprising a binder comprising at least one of polyurethane claim 1 , hydroxyl terminated polybutadiene claim 1 , or hydroxyl terminated polyether.8. The propellant of claim 7 , comprising 3% to 14% by weight the binder.9. The propellant of claim 7 , further comprising a plasticizer comprising at least one of trimethyloethane trinitrate claim 7 , triethyleneglycol dinitrate claim 7 , butannetriol trinitrate claim 7 , diethylene glycol ...

Hybrid metal composite structures, rocket motors and multi-stage rocket motor assemblies including hybrid metal composite structures, and related methods

A hybrid metal composite (HMC) structure comprises a first tier comprising a first fiber composite material structure, a second tier longitudinally adjacent the first tier and comprising a first metallic structure and a second fiber composite material structure laterally adjacent the first metallic structure, a third tier longitudinally adjacent the second tier and comprising a third fiber composite material structure, and a fourth tier longitudinally adjacent the third tier and comprising a second metallic structure and a fourth fiber composite material structure laterally adjacent the second metallic structure. At least one lateral end of the second metallic structure is laterally offset from at least one lateral end of the first metallic structure most proximate thereto. Methods of forming an HMC structure, and related rocket motors and multi-stage rocket motor assemblies are also disclosed.

SOLID PROPELLANT ROCKET MOTOR

A solid propellant rocket motor () has a tubular casing () accommodating a mass () of solid propellant material and at least one opening () for the space in the casing () to communicate with the outside closed by a closing head (); the closing head () being coupled to the casing () by means of one or the other of two blocking portions () (A) with different strength both carried by a movement device () which can be elastically deformed and operated from the outside. 110-. (canceled)11. A solid propellant rocket motor , comprising:a cylindrical tubular outer casing accommodating a mass of solid propellant material, the cylindrical tubular outer casing having at least one opening for a space in the cylindrical tubular outer casing to communicate with an outside;a closing head for closing said at least one opening; and an annular abutting surface carried by said cylindrical tubular outer casing;', 'a first blocking portion;', 'at least a second blocking portion having a lower strength than said first blocking portion;', 'wherein said second blocking portion protrudes from said first blocking portion, and said first and second blocking portions form part of a radially elastically deformable, annular open body made in one piece;', 'wherein said annular open body include opposite end portions adjacent to each other; and', 'support and moving means of said first and second blocking portions, wherein said support and moving means include actuator spacing means interposed between said end portions for spacing in a circumferential direction the end portions relative to each other and for elastically deforming said annular open body., 'means for connecting said closing head to a connection portion of said cylindrical tubular casing; said means for connecting include12. The solid propellant rocket motor according to claim 11 , wherein said first blocking portion includes a curved elongated portion comprising a plurality of said second blocking portions distributed along an outer ...

COMBUSTION CHAMBER LINER WITH SPIRAL COOLING CHANNELS

A combustion chamber liner comprising a plurality of spiral cooling channels is formed from a material component. A combustion chamber liner body extends from a first end and a second end. The combustion chamber liner body comprises a combustion chamber liner internal wall that defines a combustion area cavity extending from the first and second end. The combustion chamber liner body also comprises a combustion chamber liner external wall that is opposite the internal wall. The combustion chamber liner body also defines an inlet port, a nozzle exit port opposite the inlet port, and a throat portion. Along the combustion chamber liner external wall, spiral cooling channels are cut into the external wall such that the spiral cooling channels extend between the first and second end. 1. A combustion chamber liner for a rocket propulsion system , comprising:a combustion chamber liner body extending between a first and second end;a combustion chamber liner internal wall defined by the combustion chamber liner body, the combustion chamber liner internal wall defining a combustion area cavity extending between the first and second end;a combustion chamber liner external wall defined by the combustion chamber liner body, the combustion chamber liner external wall opposite the internal wall;an inlet port defined by the combustion chamber liner body at the first end;a nozzle exit port defined by the combustion chamber liner body at the second end;a throat portion defined by a portion of the combustion chamber liner body between the first and second end; anda plurality of spiral cooling channels defined by the combustion chamber liner external wall, the plurality of spiral cooling channels extending between the first and second end.2. The combustion chamber liner of claim 1 , wherein each spiral cooling channel of the plurality of spiral cooling channels comprises a first chamfer at a first cooling channel end and a second chamfer at a second cooling channel end.3. The ...

Variable thrust catapult

A rocket catapult assembly for an ejection seat may comprise a drive motor, a metering tube, a first cartridge, and a second cartridge. The metering tube may include an outer wall having a gas pervious section and a gas impervious section. The drive motor may be configured to translate the metering tube and align the gas pervious section or gas impervious section with a first cartridge and a second cartridge to produce a desired thrust of the rocket catapult assembly.

ROCKET ENGINE INCLUDING A LASHING DEVICE

A rocket engine benefiting from better behavior during its starting stage, the rocket engine () including a diverging section () and a lashing system () configured to hold the diverging section () while starting the rocket engine (), the lashing system () comprising: a plurality of radial cables () connected at respective first ends to a plurality of points of the diverging section (), and a peripheral cable () connected to the second ends of the radial cables () and configured to co-operate with attachment points () of a launch platform (). 1. A rocket engine including a diverging section and a lashing system configured to hold the diverging section while starting the rocket engine , the lashing system comprising:a plurality of radial cables connected at respective first ends to a plurality of points of the diverging section; anda peripheral cable connected to the second ends of the radial cables and configured to co-operate with attachment points of a launch platform.2. The rocket engine according to claim 1 , wherein the points where the radial cables are connected to the diverging section are situated in such a manner that their center of gravity lies on the main axis of the diverging section.3. The rocket engine according to claim 1 , wherein the points where the radial cables are connected to the diverging section are spaced apart in regular manner around the axis of the diverging section.4. The rocket engine according to claim 1 , wherein at least one radial cable is slidable relative to the peripheral cable.5. The rocket engine according to claim 1 , wherein the peripheral cable includes a tensioner.6. The rocket engine according to claim 5 , wherein the tensioner includes a damper.7. The rocket engine according to claim 1 , wherein the peripheral cable is configured to break in less than 5 seconds on being exposed to a temperature higher than 2500° C.8. The rocket engine according to claim 1 , wherein the peripheral cable includes a breakable segment ...

ROTATING DETONATION ACTUATOR

A flow control system includes at least one flow surface; and at least one rotating detonation actuator including: an annulus extending from an inlet end to an outlet end; an inner wall defining a radially inner boundary of the annulus; and an outer wall defining a radially outer boundary of the annulus. At least one rotating detonation wave travels through the annulus from the inlet end to the outlet end. Combustion gas from the at least one rotating detonation actuator modifies at least one flow characteristic at the flow surface. 1. A flow control system comprising:at least one flow surface; and an annulus extending from an inlet end to an outlet end;', 'an inner wall defining a radially inner boundary of the annulus; and', 'an outer wall defining a radially outer boundary of the annulus,, 'at least one rotating detonation actuator comprisingwherein at least one rotating detonation wave travels through the annulus from the inlet end to the outlet end, andwherein combustion gas from the at least one rotating detonation actuator modifies at least one flow characteristic at the at least one flow surface.2. The flow control system of claim 1 , further comprising at least one radial exit disposed in the outer wall claim 1 , the at least one radial exit fluidly coupling the annulus to an exterior of the at least one rotating detonation actuator.3. The flow control system of claim 2 , the at least one radial exit further comprising multiple radial exits claim 2 ,wherein a first plurality of radial exits of the multiple radial exits is arranged in a first row of radial exits circumferentially spaced around the at least one rotating detonation actuator,wherein a second plurality of radial exits of the multiple radial exits is arranged in a second row of radial exits circumferentially spaced around the at least one rotating detonation actuator, andwherein the second row of radial exits is axially aft of the first row of radial exits.4. The flow control system of claim 1 , ...

Air-breathing rocket engine

An air-breathing rocket engine in certain embodiments comprises an outer shell and an interior portion situated entirely within the front end of the outer shell. The interior portion includes a funnel-shaped intake and an annular primary combustion chamber between the inner front wall of the shell and the outer surface of the funnel-shaped intake. The intake has a central aperture that is in fluid communication with the throat and exhaust areas within the outer shell. A second circumferential gap is formed between the outer surface of the front inner wall and the inner surface of the front end of the outer shell and is in fluid communication with the throat and exhaust areas within the outer shell. One or more injector ports and one or more ignition ports are situated at the front end of the second circumferential gap.

IMPREGNATED FIBERS COMPRISING PRECERAMIC RESIN FORMULATIONS, AND RELATED COMPOSITE MATERIALS AND METHODS

A preceramic resin formulation comprising a polycarbosilane preceramic polymer, an organically modified silicon dioxide preceramic polymer, and, optionally, at least one filler. The preceramic resin formulation is formulated to exhibit a viscosity of from about 1,000 cP at about 25° C. to about 5,000 cP at a temperature of about 25° C. The at least one filler comprises first particles having an average mean diameter of less than about 1.0 μm and second particles having an average mean diameter of from about 1.5 μm to about 5 μm. Impregnated fibers comprising the preceramic resin formulation are also disclosed, as is a composite material comprising a reaction product of the polycarbosilane preceramic polymer, organically modified silicon dioxide preceramic polymer, and the at least one filler. Methods of forming a ceramic matrix composite are also disclosed. 1. Impregnated fibers comprising fibers and a preceramic resin formulation comprising a polycarbosilane preceramic polymer , an organically modified silicon dioxide preceramic polymer , and at least one filler , the at least one filler comprising first particles having an average mean diameter of less than about 1.0 μm and second particles having an average mean diameter of from about 1.5 μm to about 5 μm.2. The impregnated fibers of claim 1 , wherein the fibers comprise polyacrylonitrile-based fibers.3. The impregnated fibers of claim 1 , wherein the fibers comprise pitch-based fibers.4. A composite material comprising fibers and a reaction product of a polycarbosilane preceramic polymer claim 1 , an organically modified silicon dioxide preceramic polymer claim 1 , and at least one filler claim 1 , the at least one filler comprising first particles having an average mean diameter of less than about 1.0 μm and second particles having an average mean diameter of from about 1.5 μm to about 5 μm.5. The composite material of claim 4 , wherein the composite material is configured as at least a portion of a rocket ...

LIQUID ROCKET ENGINE ASSEMBLIES AND RELATED METHODS

A liquid rocket engine assembly comprising a thrust chamber, a nozzle, and a joint structure. The joint structure attaches the thrust chamber and the nozzle and comprises at least one seal element and an attachment ring interposed between the thrust chamber and the nozzle. Fasteners extend between the nozzle and the thrust chamber through the at least one seal element and the attachment ring. Materials of the thrust chamber and of the nozzle comprise different coefficients of thermal expansion. A method of forming a liquid rocket engine assembly is also disclosed. 1. A joint structure for attaching a thrust chamber and a nozzle of a rocket engine assembly , the joint structure comprising:two seal elements configured to be interposed between the thrust chamber and the nozzle;an insulation ring configured to be disposed between the two seal elements;an attachment ring configured to be disposed circumferentially around the thrust chamber and adjacent to an aft of the insulation ring; andfasteners configured to extend through the attachment ring and at least one seal element of the two seal elements and to secure the joint structure to the thrust chamber and to the nozzle.2. The joint structure of claim 1 , wherein at least one seal element of the two seal elements or the attachment ring comprises an annular shape.3. The joint structure of claim 1 , wherein at least one seal element of the two seal elements comprises a flexible material.4. The joint structure of claim 1 , wherein the fasteners are configured to extend through both seal elements of the two seal elements.5. The joint structure of claim 1 , wherein the attachment ring is separate and discrete from the nozzle claim 1 , and wherein the flange element forms an integral part of the thrust chamber.6. The joint structure of claim 1 , further comprising a support ring laterally adjacent to the insulation ring and disposed at least partially between the two seal elements.7. The joint structure of claim 1 , further ...

PRECURSOR COMPOSITIONS FOR AN INSULATION, INSULATED ROCKET MOTORS, AND RELATED METHODS

An insulation precursor composition comprises ethylene propylene diene monomer, an aramid, and a bromine-containing flame retardant. Rocket motors comprising a case, an energetic material within the case, and an insulation material comprising a reaction produce of ethylene propylene diene monomer, an aramid, and a flame retardant comprising bromine are also disclosed. Related precursor compositions are also disclosed. 1. An insulation precursor composition , comprising:ethylene propylene diene monomer;an aramid; anda bromine-containing flame retardant.2. The insulation precursor composition of claim 1 , further comprising antimony trioxide.3. The insulation precursor composition of claim 2 , wherein the insulation precursor composition comprises from about 4 moles to about 5 moles of the antimony trioxide for every about 1 mole of bromine in the bromine-containing flame retardant.4. The insulation precursor composition of claim 1 , wherein the bromine-containing flame retardant further comprises chlorine.5. The insulation precursor composition of claim 1 , wherein the bromine-containing flame retardant comprises decabromodiphenyl ethane.6. The insulation precursor composition of claim 1 , wherein the bromine-containing flame retardant comprises ethylenebistetrabromophthalimide.7. The insulation precursor composition of claim 1 , wherein the bromine-containing flame retardant comprises one or more of tris(2 claim 1 ,3-dibromoisopropyl) isocyanurate claim 1 , tetrabromobisphenol A claim 1 , tetrabromobisphenol A bis(2 claim 1 ,3-dibromopropyl ether) claim 1 , and a brominated styrene.8. The insulation precursor composition of claim 1 , wherein the bromine-containing flame retardant constitutes from about 15 weight percent to about 40 weight percent of the insulation precursor composition.9. The insulation precursor composition of claim 1 , further comprising polybutadiene.10. The insulation precursor composition of claim 1 , further comprising a curing agent.11. The ...

Device for Burning Off Propellants Or Explosive Substances

The invention relates to a device for burning off propellants or explosive substances, which has an activation temperature that lies below the spontaneous ignition temperature of the propellant or explosive substance. The device () comprises at least two substances () reacting exothermically with one another, wherein at least one first substance () is present in a liquid aggregate state below the activation temperature of the device and is separated from at least one second substance () by at least one pressure-tight barrier (). 1. A device for burning fuels or explosives at a trigger temperature that is below the autoignition temperature of the fuel or explosive , the device comprising:at least two substances including a first substance and a second substance that react exothermically with each other,wherein the first substances exists in a liquid state of aggregation when below the trigger temperature of the device and is separated from the second substance by at least one pressure-tight barrier,wherein, upon reaching the trigger temperature, the pressure-tight barrier allows contact between the first substance and the second substance.2. The device according to claim 1 , wherein the pressure-tight barrier comprises at least one material having a melting temperature at least equal to the trigger temperature of the device.3. The device according to claim 1 , wherein the pressure-tight barrier features at least one flow-through opening that is closed off with a material having a melting temperature equal to the trigger temperature of the device.4. The device according to claim 1 , wherein the pressure-tight barrier is made from at least one membrane comprising a puncture mechanism.5. The device according to claim 1 , wherein the pressure-tight barrier takes the form of a seal that features at least one flow-through opening claim 1 , where the flow-through opening is closed off by a piston until the trigger temperature of the device is reached.6. The device according ...

Multi-Stage Solid Rocket Motor

A multi-stage solid rocket in accordance with this disclosure includes a first or primary solid fuel core that operates as a standard solid rocket. A secondary solid fuel core may be “wrapped around” the primary solid fuel core in a layered arrangement so as to use the same casing. The secondary solid fuel core can be configured with an insufficient amount of oxidizer to burn by itself or to be ignited by the primary solid fuel core during its burn. The oxidizer necessary to enable a secondary solid fuel core burn can be controllably released from a secondary source to permit variable thrust generation. Subsequent cores may be wrapped around prior cores and configured with insufficient amounts of oxidizer to be ignited by any prior core. The oxidizer necessary to enable any subsequent core burn may also be controllably released to permit variable thrust generation. 1. A multi-stage solid rocket , comprising:a structural casing;a first solid fuel core enclosed in the structural casing and having first fuel and first oxidizer, wherein the first oxidizer is configured to supply sufficient oxygen to sustain combustion of the first fuel;a second solid fuel core enclosed in the structural casing and in a layered configuration with the first solid fuel core, the second solid fuel core having second fuel and second oxidizer, wherein the second oxidizer is configured to supply insufficient oxygen to sustain combustion of the second fuel, and wherein the first oxidizer is further configured to provide insufficient oxygen to ignite or sustain combustion of the second solid fuel core; anda second oxygen source separate from, and in fluid communication with, the second solid fuel core, wherein the second oxygen source is configured to supply sufficient oxygen to, in combination with the second oxidizer, maintain combustion of the second fuel.2. The multi-stage solid rocket of claim 1 , further comprising a first igniter system configured to ignite the first solid fuel core.3. ...

PRECURSOR COMPOSITIONS FOR AN INSULATION, INSULATED ROCKET MOTORS, AND RELATED METHODS

A precursor composition comprising, before curing, ethylene propylene diene monomer (EPDM), silica, magnesium hydroxide, polymerized 1,2-dihydro-2,2,4-trimethylquinoline, a solid chlorinated paraffin, stearic acid, a five carbon petroleum hydrocarbon, trimethylolpropane trimethacrylate, and a peroxide. A rocket motor including a reaction product of the precursor composition and a method of insulating a rocket motor. 1. A precursor composition , comprising , before curing:ethylene propylene diene monomer (EPDM), silica, magnesium hydroxide, polymerized 1,2-dihydro-2,2,4-trimethylquinoline, a solid chlorinated paraffin, stearic acid, a five carbon petroleum hydrocarbon, a co-agent, and a peroxide.2. The precursor composition of claim 1 , wherein the EPDM comprises a terpolymer of ethylene claim 1 , propylene claim 1 , and ethylene norbornene.3. The precursor composition of claim 1 , wherein the EPDM comprises an ethylene content of about 53% by weight and a diene content of about 6.0% by weight.4. The precursor composition of claim 1 , wherein the EPDM comprises controlled long chain branching.5. The precursor composition of claim 1 , wherein the silica comprises amorphous claim 1 , precipitated silica.6. The precursor composition of claim 1 , wherein the peroxide comprises 1 claim 1 ,1-di-(t-butylperoxy)-3 claim 1 ,3 claim 1 ,5-trimethylcyclohexane or dicumyl peroxide.7. The precursor composition of claim 1 , wherein the co-agent comprises trimethylolpropane trimethacrylate or a poly(butadiene) resin.8. The precursor composition of claim 1 , wherein the EPDM comprises from about 70 parts to about 150 parts of the precursor composition claim 1 , the silica comprises from about 21 parts to about 45 parts of the precursor composition claim 1 , the magnesium hydroxide comprises from about 1.0 part to about 7.0 parts claim 1 , the polymerized 1 claim 1 ,2-dihydro-2 claim 1 ,2 claim 1 ,4-trimethylquinoline comprises from about 0.35 part to about 0.75 part of the precursor ...

PRECURSOR COMPOSITIONS FOR AN INSULATION, INSULATED ROCKET MOTORS, AND RELATED METHODS

A precursor composition comprising, before curing, ethylene propylene diene monomer (EPDM), zinc oxide, silica, polymerized 1,2-dihydro-2,2,4-trimethylquinoline, a solid chlorinated paraffin, stearic acid, a five carbon petroleum hydrocarbon, trimethylolpropane trimethacrylate, and a peroxide. A rocket motor including a reaction product of the precursor composition and a method of insulating a rocket motor. 1. A precursor composition , comprising , before curing:ethylene propylene diene monomer (EPDM), zinc oxide, silica, polymerized 1,2-dihydro-2,2,4-trimethylquinoline, a solid chlorinated paraffin, stearic acid, a five carbon petroleum hydrocarbon, trimethylolpropane trimethacrylate, and a peroxide.2. The precursor composition of claim 1 , wherein the EPDM comprises a terpolymer of ethylene claim 1 , propylene claim 1 , and ethylene norbornene.3. The precursor composition of claim 1 , wherein the EPDM comprises an ethylene content of about 50.0% by weight and a diene content of about 5.0% by weight.4. The precursor composition of claim 1 , wherein the zinc oxide comprises a propionic acid coated zinc oxide.5. The precursor composition of claim 1 , wherein the silica comprises amorphous claim 1 , precipitated silica.6. The precursor composition of claim 1 , wherein the peroxide comprises 1 claim 1 ,1-di-(t-butylperoxy)-3 claim 1 ,3 claim 1 ,5-trimethylcyclohexane.7. The precursor composition of claim 1 , wherein the EPDM comprises from about 70 parts to about 150 parts of the precursor composition claim 1 , the zinc oxide comprises from about 2.1 parts to about 4.5 parts of the precursor composition claim 1 , the silica comprises from about 21 parts to about 45 parts of the precursor composition claim 1 , the polymerized 1 claim 1 ,2-dihydro-2 claim 1 ,2 claim 1 ,4-trimethylquinoline comprises from about 0.35 part to about 0.75 part of the precursor composition claim 1 , the solid chlorinated paraffin comprises from about 4.2 parts to about 9 parts of the precursor ...

HYBRID METAL COMPOSITE STRUCTURES, ROCKET MOTORS INCLUDING HYBRID METAL COMPOSITE STRUCTURES, AND RELATED METHODS

A hybrid metal composite (HMC) structure comprises a first tier comprising a first fiber composite material structure, a second tier longitudinally adjacent the first tier and comprising a first metallic structure and a second fiber composite material structure laterally adjacent the first metallic structure, a third tier longitudinally adjacent the second tier and comprising a third fiber composite material structure, and a fourth tier longitudinally adjacent the third tier and comprising a second metallic structure and a fourth fiber composite material structure laterally adjacent the second metallic structure. At least one lateral end of the second metallic structure is laterally offset from at least one lateral end of the first metallic structure most proximate thereto. Methods of forming an HMC structure, and related rocket motors and multi-stage rocket motor assemblies are also disclosed. 1. A hybrid metal composite (HMC) structure , comprising stacked tiers comprising: a fiber platform comprising a single, wound tow comprising fibers substantially aligned with one another in a single direction; and', 'a matrix material surrounding the fiber platform; and, 'first tiers comprising a fiber composite material, within each of the first tiers the fiber composite material comprisingsecond tiers longitudinally alternating with the first tiers and comprising metallic structures, the fiber composite material directly adjacent opposing lateral sides of the metallic structures, and the opposing lateral sides of each of the metallic structures laterally offset from corresponding ends of a longitudinally neighboring one of the metallic structures most longitudinally proximate thereto.2. The HMC structure of claim 1 , further comprising:metal plates overlying and underlying the stacked tiers; andone or more fasteners longitudinally extending through each of the metal plates, the metallic structures, and the fiber composite material, the one or more fasteners comprising a ...

Thruster with segmented propellant

A thruster includes multiple segments of electrically-operated propellant, electrodes for igniting one or a few of the electrically-operated propellant segments at a time, and a propellant feeder for moving further propellant segments into engagement with the electrodes. The segments may be configured to provide equal increments of thrust, or different amounts of thrust. The segments may each include an electrically-operated propellant material surrounded by a sealing material, so as to keep the propellant material away from moisture and other contaminants (and/or the vacuum of space) before each individual segment is to be used. The thruster may be included in any of a variety of flight vehicles, for example in a small satellite such as a CubeSat satellite, for instance having a volume of about 1 liter, and a mass of no more than about 1.33 kg.

AUXILIARY BOOSTER WITH OPTIMISED ARCHITECTURE

A solid propellant auxiliary booster intended to be attached to the main body of a launcher comprises a cylindrical body extending in a longitudinal direction between a rear face in communication with a nozzle and a front face formed by a conical structure connected to the cylindrical body of the booster. The cylindrical body delimits a first internal volume and the conical structure of the front face delimits a second internal volume. The auxiliary booster contains a solid propellant charge. The first internal volume of the cylindrical body communicates with the second internal volume of the conical structure. The solid propellant charge is present both in the first and second internal volumes.

HYBRID ROCKET ENGINE WITH IMPROVED SOLID FUEL SEGMENT

A rocket engine with an improved solid fuel segment mainly comprises a combustion chamber, a solid fuel segment installed in the combustion chamber, and an oxidizer injector installed at one end of the combustion chamber. The solid fuel segment surrounds and forms a trajectory to allow the oxidizer injector to inject oxidizer into the trajectory, in particular, on the solid fuel segment is formed with a plurality of protrusions, between the each two protrusions are defined a recess, a flame holding hot-gas region is formed between the protrusion and the recess, so as to produce eddy current in the flame holding hot-gas region when the propellant mixture is burned inside the trajectory, such that the whole solid fuel segment can produce even regression rate and high combustion efficiency.

RIGID THERMAL PROTECTION COMPOSITION

A polymer composite composition for use in high temperature applications such as furnaces, heat shields and aeronautical jet and rocket motors. In a particular application, the disclosed composition is applied to the manufacture of rocket motor cases, or parts thereof, to provide rigid thermal protection (RTP). The polymer composite composition comprises cyanate ester resin, fine lengths of carbon fibre and refractory filler material. 1. A mouldable polymer composite composition comprising:(i) cyanate ester resin;(ii) fine lengths of carbon fibre; and(iii) refractory filler material.2. The composition of claim 1 , wherein components (i) claim 1 , (ii) and (iii) are present in a weight ratio of 75-100:5-50:5-50 (resin:carbon fibre:filler).3. The composition of claim 1 , wherein components (i) claim 1 , (ii) and (iii) are present in a weight ratio of 90-100:10-30:10-30 (resin:carbon fibre:filler).4. The composition of claim 1 , wherein components (i) claim 1 , (ii) and (iii) are present in a weight ratio of 90-100:5-15:10-35 (resin:carbon fibre:filler).5. The composition of claim 1 , wherein the composition is compression mouldable.6. The composition of claim 1 , wherein the cyanate ester resin is a blend of PT-30 and LeCy resins.7. The composition of claim 1 , wherein the carbon fibre is present in the composition in a form of fine fibre lengths wherein individual fibre lengths vary in length from about 3 mm to about 30 mm.8. The composition of claim 1 , wherein the refractory filler material comprises silica.9. The composition of claim 1 , wherein the refractory filler material comprises pyrogenic silica comprising hydrophilic and/or amorphous particles that claim 1 , substantially claim 1 , are less than 50 μm.10. The composition of claim 1 , wherein components (i) claim 1 , (ii) and (iii) together comprise at least 90% by weight of the total weight of the composition.11. The composition of claim 10 , wherein by itself claim 10 , the cyanate ester resin (component ...

INNER COATING LAYER FOR SOLID-PROPELLANT ROCKET ENGINES

An inner coating layer for solid-propellant rocket engines, constituted by a material comprising from 45% to 55% wt. of a a cross-linkable, unsaturated-chain polymer base, from 11% to 13% wt. of silica, from 15% to 25% wt. of vulcanizing agents and plasticizers, from 5% to 7% wt. of aramid fiber and from 10% to 15% wt. of microspheres made of a material selected among glass, quartz and nano clay, having diameter lower than 200 gm, density comprised between 0.30 and 0.34 g/cc and resistance to hydrostatic pressure greater than, or equal to, 4500 psi. 1. An inner coating layer for solid-propellant rocket engines , the inner coating comprising:an ablative material comprising from 45% to 55% wt. of a cross-linkable, unsaturated-chain polymer base, from 11% to 13% wt. of silica and from 15% to 25% wt. of vulcanizing agents and plasticizers, the ablative material further comprising from 5% to 7% wt. of aramid fiber and from 10% to 15% wt. of micro-spheres made of a material selected among glass, quartz and nano clay, the micro-spheres having a diameter lower than 200 um, the micro-spheres having a density comprised between 0.30 and 0.34 g/cc and the micro-spheres having a resistance to hydrostatic pressure greater than or equal to 4500 psi.2. The inner coating layer of claim 1 , wherein the micro-spheres have conductivity comprised between 0.07 and 0.22 W/m° K.3. The inner coating layer of claim 1 , wherein the polymer base is comprised in the group constituted by EPDM claim 1 , NBR claim 1 , SBR claim 1 , HTPB.4. The inner coating layer of claim 1 , wherein the aramid fiber is Kevlar® or Twaron®.5. The inner coating layer of claim 1 , wherein the aramid fiber is added to a mixture in dried form.6. The inner coating layer of claim 2 , wherein the polymer base is at least one of EPDM claim 2 , NBR claim 2 , SBR and HTPB.7. The inner coating layer of claim 2 , wherein the aramid fiber is one of Kevlar® and Twaron®.8. The inner coating layer of claim 3 , wherein the aramid ...

In-situ solid rocket motor propellant grain aging using gas

A method for non-destructively determining a mechanical property of a solid rocket motor propellant grain may comprise applying, via a gas, a force to a surface of the solid rocket motor propellant grain, wherein a deformation is formed on the surface of the solid rocket motor propellant grain in response to the applying, and measuring a pressure of the gas. This process may be performed over time to determine a lifespan of the propellant grain.

Rocket Motors and Their Use

Apparatus, methods and computer programs are provided. In one example, an apparatus is a rocket motor, comprising: a casing having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension; and propellant, located inside the casing, arranged to generate a force in a direction that is substantially perpendicular to the length dimension of the casing. 1. A rocket motor , comprising:a casing having a length dimension, a width dimension and a depth dimension, wherein the length dimension is greater than the width dimension and greater than the depth dimension; andpropellant, located inside the casing, arranged to generate a force in a direction that is substantially perpendicular to the length dimension of the casing.2. The rocket motor as claimed in claim 1 , wherein the propellant is arranged to cause ejection of gas from the casing in a direction that is substantially aligned with the depth dimension of the casing.3. The rocket motor as claimed in claim 1 , wherein the propellant is arranged to cause ejection of gas from the casing in a direction that is substantially perpendicular to the length dimension of the casing.4. The rocket motor as claimed in claim 1 , wherein the length dimension is orthogonal to the width dimension claim 1 , the length dimension is orthogonal to the depth dimension claim 1 , and the width dimension is orthogonal to the depth dimension.5. The rocket motor as claimed in claim 1 , wherein the propellant is in the form of pellets.6. The rocket motor as claimed in claim 1 , wherein the casing comprises a plurality of exit gas apertures.7. The rocket motor as claimed in claim 6 , wherein at least some of the gas exit apertures diverge in the direction of movement of gas ejected from the casing in operation.8. The rocket motor as claimed in claim 1 , wherein the casing defines one or more rocket nozzles.9. The rocket motor as claimed in claim ...

INTEGRATED ROCKET MOTOR AGING SENSOR

A solid rocket motor propellant grain arrangement may comprise a case, a propellant grain disposed within the case, and an integrated rocket motor aging sensor disposed outward from the propellant grain, wherein the integrated rocket motor aging sensor is configured to measure data corresponding to a plurality of distinct locations of the propellant grain. The integrated rocket motor aging sensor may comprise a resistive screen matrix (RSM). 1. A method for non-destructively determining a health of a solid rocket motor propellant grain , comprising:receiving, by a controller, a first data corresponding to a plurality of distinct locations of the solid rocket motor propellant grain from an integrated rocket motor aging sensor at a first time;receiving, by the controller, a second data corresponding to the plurality of distinct locations of the solid rocket motor propellant grain from the integrated rocket motor aging sensor at a second time; andcomparing, by the controller, the first data with the second data.2. The method of claim 1 , wherein the integrated rocket motor aging sensor comprises a resistive screen matrix (RSM).3. The method of claim 2 , wherein the second data indicates at least one of an expansion or contraction of the solid rocket motor propellant grain.4. The method of claim 2 , wherein the solid rocket motor propellant grain is a solid mass with an exposed inner surface area defining a perforation volume in the interior of the solid rocket motor propellant grain.5. The method of claim 2 , wherein the first data corresponds to the plurality of distinct locations of an outer surface of the solid rocket motor propellant grain.6. The method of claim 2 , wherein receiving the first data comprises receiving a plurality of first datum corresponding to a plurality of nodes of the RSM claim 2 , wherein each node corresponds to at least one of the plurality of distinct locations.7. The method of claim 6 , wherein receiving the second data comprises receiving ...

Solid fuel gas generator

FIELD: machine engineering. SUBSTANCE: gas generator comprises a central hollow cylinder closed at one end and open in the form of a converging nozzle from the other end, local gas-generating parts of solid fuel charge located in the cylinder with separating partitions therebetween, and an ignition system of the charge parts. The gas generator further comprises a cylindrical body with peripheral and central filler charges, a resonance chamber and two circular grids. The cylindrical body is provided with a bottom on one side and a conical adapter with a branch pipe on the other side. The resonance chamber is made in the form of a cup with a piston located therein and installed in the center of the body bottom. The filler charges are installed in the body between inlet and outlet circular grids with formation of a circular channel therebetween and coaxial the hollow cylinder. The hollow cylinder is in axial hole of the central filler charge and faces the nozzle along the axis towards the resonance chamber with formation of an intermediate cavity therebetween communicating through the circular channel between the filler charges with the branch pipe. EFFECT: possibility for repeated activation of the gas generator and controlling the mass flow rate of gaseous low-temperature working medium, increased reliability of the gas generator. 5 cl, 3 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 633 976 C1 (51) МПК F23R 5/00 (2006.01) F02K 9/32 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (21)(22) Заявка: 2016122675, 09.06.2016 (24) Дата начала отсчета срока действия патента: 09.06.2016 Дата регистрации: (73) Патентообладатель(и): Федеральное государственное унитарное предприятие "Центральный институт авиационного моторостроения имени П.И. Баранова" (RU) Приоритет(ы): (22) Дата подачи заявки: 09.06.2016 (45) Опубликовано: 20.10.2017 Бюл. № 29 Стр.: 1 C 1 2 6 3 3 9 7 6 R U 143274 U1, 20.07.2014. RU 2131053 C1, 27.05.1999. RU ...

Jet ammunition engine

FIELD: military equipment. SUBSTANCE: invention relates to jet engines for ammunition intended for firing from anti-sabotage grenade launcher systems, arranged on floating warships, stationary or mobile objects on shore. Jet propellant engine comprises combustion chamber 2 with tubular powder charge 3 of "brush" structure arranged therein, separated from igniter 5 by diaphragm 4, having through holes and nozzle. Separating diaphragm in the chamber is installed movably, and on its external contour there are evenly located holes, which form along the perimeter of the diaphragm projections in the form of teeth 10, wherein the ratio of the total area of the holes to the area of the circle described by the teeth ranges from 0.279 to 0.339, and the ratio of the outer diameter of the diaphragm to the inner diameter of the combustion chamber ranges from 0.997 to 1.028. Separation diaphragm is made from non-metallic material with thickness of at least 1 mm, for example, cardboard. EFFECT: invention ensures preservation of igniter at transport loads and simplifies diaphragm design. 3 cl, 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 748 027 C2 (51) МПК F02K 9/95 (2006.01) F02K 9/32 (2006.01) F02K 9/36 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/95 (2020.08); F02K 9/32 (2020.08); F02K 9/36 (2020.08) (21)(22) Заявка: 2019126088, 16.08.2019 (24) Дата начала отсчета срока действия патента: Дата регистрации: 19.05.2021 (43) Дата публикации заявки: 16.02.2021 Бюл. № 5 2 7 4 8 0 2 7 R U Адрес для переписки: 119160, Москва, Фрунзенская наб., 22/2, Управление интеллектуальной собственности, военно-технического сотрудничества и экспертизы поставок вооружения и военной техники Министерства обороны Российской Федерации (54) ДВИГАТЕЛЬ РЕАКТИВНОГО БОЕПРИПАСА (57) Реферат: Изобретение относится к области военной техники, в частности, к реактивным двигателям для боеприпасов, предназначенных для стрельбы из ...

Systems, devices and methods of ignition suppression

FIELD: electrical engineering. SUBSTANCE: invention relates to electrical engineering and fire extinguishing and can be used in operation of fuel systems of vehicles to prevent occurrence of various inflammations. Proposed anti-inflammable system comprises fire-hazardous structure, which passes from support structure into combustible medium, and contains a porous anti-inflammable coating which substantially covers a fire-hazardous structure, wherein the anti-inflammable coating is configured to suppress the ignition phenomenon caused by the ignition source associated with the fire-hazardous structure, and comprises a porous body which can contain one or more porous elements. Proposed ignition suppression method includes installation of anti-inflammable coating over fire-hazardous structure to prevent volumetric combustion, for example, fuel vapours in a fuel tank due to an ignition event associated with a fire-hazardous structure. EFFECT: technical result of the invention is reduced probability of ignition caused by ignition source, as well as prevention of volumetric combustion, for example of fuel vapours. 20 cl, 5 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 700 670 C2 (51) МПК F02D 41/22 (2006.01) A62C 2/06 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02D 41/22 (2019.05); A62C 2/06 (2019.05) (21)(22) Заявка: 2016114320, 13.04.2016 (24) Дата начала отсчета срока действия патента: Дата регистрации: 20.09.2019 (73) Патентообладатель(и): Зе Боинг Компани (US) 21.07.2015 US 14/805,259 (43) Дата публикации заявки: 18.10.2017 Бюл. № 29 (45) Опубликовано: 20.09.2019 Бюл. № 26 2 7 0 0 6 7 0 (54) СИСТЕМЫ, УСТРОЙСТВА И СПОСОБЫ ПОДАВЛЕНИЯ ВОСПЛАМЕНЕНИЯ (57) Реферат: Изобретение относится к области структурой, и содержит пористое тело, которое электротехники и пожаротушения и может быть может содержать один или большее количество использовано при эксплуатации топливных пористых элементов. Предложенный способ ...

Charge rocket engine for de-mining charge

FIELD: machine building.SUBSTANCE: invention relates to solid-propellant rocket engines used for air supply of demining charge at preset distance when engine is used in mine-clearing installations. Rocket engine includes shell, front cover, nozzle unit with six nozzles inclined to engine longitudinal axis, igniter and single-channel charge in the form of solid propellant rock cartridge armored along outer surface and ends. Two groups of internal radial grooves are made at the front and nozzle end faces of the charge so that grooves of each group have different distance to external surface of charge fuel. For each group of grooves distance from charge end to groove, closest to end, is 0.8–1 distance from groove to external surface of fuel charge. Distance between adjacent surfaces of adjacent grooves makes 1.6–2 of distance from groove to external surface of charge fuel. Charge is arranged between the front and nozzle supports, and the nozzle support is equipped with a shutter.EFFECT: invention allows improving operating efficiency of rocket engine, as well as its ignition reliability.1 cl, 5 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 711 328 C1 (51) МПК F02K 9/18 (2006.01) F02K 9/30 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/18 (2019.08); F02K 9/30 (2019.08) (21)(22) Заявка: 2018140228, 15.11.2018 (24) Дата начала отсчета срока действия патента: Дата регистрации: 16.01.2020 (45) Опубликовано: 16.01.2020 Бюл. № 2 Адрес для переписки: 140090, Московская обл., г. Дзержинский, ул. Академика Жукова, 42, ФГУП "ФЦДТ "Союз" C 1 2 7 1 1 3 2 8 R U (56) Список документов, цитированных в отчете о поиске: US 5385099 A, 31.01.1995. US 4052943 A, 11.10.1977. RU 2326260 C2, 10.06.2008. US 2002/0062756 A1, 30.05.2002. RU 2298110 C2, 27.04.2007. (54) РАКЕТНЫЙ ДВИГАТЕЛЬ ПОДАЧИ ЗАРЯДА РАЗМИНИРОВАНИЯ (57) Реферат: Изобретение относится к ракетным наружной поверхности и торцам. У переднего и двигателям твердого ...

Ракетный двигатель твердого топлива для увода отделяемых частей

Ракетный двигатель твердого топлива для увода отделяемых частей ракеты состоит из корпуса с твердотопливным многошашечным зарядом, расположенным между опорными решетками и двумя газосвязанными соплами. В предсопловых объемах корпуса соосно газоподводной трубке с пиропатроном закреплены воспламенители, каждый из которых содержит перфорированный стакан с установленным внутри футляром, заполненным пиротехническим составом. Стакан со стороны газоподводной трубки закрыт крышкой с кольцевым коническим отражателем. Между стенкой отражателя и торцом футляра в боковой стенке крышки выполнен круговой ряд сквозных каналов, соединяющих внутреннюю полость с предсопловым объемом. Наружный диаметр кольцевого конического отражателя dсоставляет 0,4…0,5 от внутреннего диаметра корпуса (d=(0,4…0,5)D). В стенке отражателя на диаметре d, равном 0,6…0,8 от наружного диаметра конического отражателя (d=(0,6… 0,8)d), дополнительно выполнен ряд сквозных отверстий, оси которых расположены перпендикулярно внешней стороне отражателя. Во фронтальном сечении оси отверстий в отражателе смещены относительно осей сквозных каналов крышки и расположены в секторе между ними, а диаметр кругового ряда отверстий в отражателе больше диаметра кругового ряда отверстий в крышке. В отражателе может быть выполнено более одного ряда отверстий. Изобретение позволит уменьшить влияние действия волн давления продуктов сгорания заряда за счет уменьшения амплитуды волн давления. 1 з.п. ф-лы, 4 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 678 602 C1 (51) МПК F02K 9/32 (2006.01) F02K 9/30 (2006.01) F02K 9/95 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/32 (2018.08); F02K 9/30 (2018.08); F02K 9/95 (2018.08) (21)(22) Заявка: 2017145529, 25.12.2017 (24) Дата начала отсчета срока действия патента: Дата регистрации: 30.01.2019 (45) Опубликовано: 30.01.2019 Бюл. № 4 (56) Список документов, цитированных в отчете о поиске: RU 2500913 C1, 10.12.2013. RU 2 6 7 ...

Cumulative-high-explosive charge engine

FIELD: engines and pumps.SUBSTANCE: invention relates to the solid fuel rocket engines used for operation as a part of cumulative high-explosive charge. Cumulative high-explosive charge engine comprises housing, nozzle, placed between the grate and the transition bottom charge, igniter and membrane in the form of cover. Grate is installed on the nozzle side and represents the series of concentric rings, between which holes and grooves are made. Membrane is mounted on the nozzle and has an afterburner tube on the engine inner side, and from the outside on the diaphragm the threaded socket for electric igniter or trigger is made. Igniter is placed between the grate and the retainer, which is fixed in the gap between the nozzle and the grate. Retainer is the convex thin-walled sheet part, in which center there is a hole, and on the retainer flange local protrusions are made by molding. Protrusions height is greater than the gap between the nozzle and the grate, and the holes in the grate orifice area ratio to the retainer hole area is greater than the charge accommodation chamber free volume ratio to the pre-nozzle volume.EFFECT: invention allows to increase the engine efficiency and ensure the charge reliable ignition.1 cl, 5 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 675 983 C1 (51) МПК F02K 9/95 (2006.01) F02K 9/36 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/95 (2018.08); F02K 9/36 (2018.08) (21)(22) Заявка: 2018106641, 22.02.2018 (24) Дата начала отсчета срока действия патента: Дата регистрации: 25.12.2018 (45) Опубликовано: 25.12.2018 Бюл. № 36 2 6 7 5 9 8 3 R U (73) Патентообладатель(и): Федеральное государственное унитарное предприятие "Федеральный центр двойных технологий "Союз" (ФГУП "ФЦДТ "Союз") (RU) (56) Список документов, цитированных в отчете о поиске: RU 2604772 C1, 10.12.2016. RU 2333379 C1, 10.09.2008. US 2979896 A, 18.04.1961. RU 2133864 C1, 27.07.1999. US 5169093 A, 08.12.1992. ...

Aramid polymer and powder filler reinforced elastomeric composition for use as a rocket motor insulation

An elastomeric composition suitable for use as a rocket motor case insulation is disclosed. The composition consists of a vulcanizable elastomeric composition and reinforcing aramid polymer fibers in combination with a powder filler. A preferred embodiment utilizes polyisoprene as the elastomer, KELVAR® fibers as the aramid polymer fibers, and silica as the powder filler.

一种适用自由装填固体火箭发动机的耐烧蚀药柱支撑板

本发明提供了一种适用自由装填固体火箭发动机的耐烧蚀药柱支撑板，包括上支撑板、下支撑板及沉头螺钉。所述上支撑板由上支撑环、包覆圆柱筋棒组成；下支撑板由下支撑环、包覆圆柱筋棒组成；上支撑板和下支撑板通过沉头螺钉连接，形成双向正交支撑结构。本发明采用的药柱支撑板结构，不仅可以实现对自由装填药柱的轴向定位和固定；还可以在固体火箭发动机工作过程中承受大秒流量高温高压燃气的烧蚀作用，避免支撑板因高温燃气烧蚀作用发生筋板断裂或出现较大面积烧熔孔导致自由装填药柱燃烧不充分或异常飞出，提高了发动机工作过程的可靠性，从而保证了发动机的工作性能，同时提高了结构完整性及装配便捷性。

Solid-propellant rocket engines

FIELD: machine building.SUBSTANCE: invention relates to rocket propulsion units, namely, to solid-propellant rocket engines with charges from mixed solid fuels with inner combustion surface of star-like or similar shape, which are firmly fixed with heat-protective coating applied to the inner surface of the housing through protective-fixing layer. Solid-propellant rocket engine includes housing, heat-protective coating laid along its inner surface, end collars, protective-fixing layer and solid fuel charge. Along the entire length of the charge between the heat-insulating coating and the protective-fixing layer there is an intermediate layer of metal foil, which is evenly attached to the heat-insulating coating and the protective-fixing layer by an adhesive bond.EFFECT: metal foil layer prevents penetration of gaseous components (ammonia) released into solid fuel charge from heat-insulating coating through protective-fixing layer during storage and operation, which prevents destruction of urethane and ester bonds of polydiene-urethane epoxy rubber, which is the base of charge of solid fuel, provides design traction characteristics during the whole operation of the solid-propellant rocket engine and improves reliability of engine operation.1 cl, 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 711 892 C1 (51) МПК F02K 9/34 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/346 (2019.08) (21)(22) Заявка: 2018144972, 19.12.2018 (24) Дата начала отсчета срока действия патента: Дата регистрации: 23.01.2020 (45) Опубликовано: 23.01.2020 Бюл. № 3 2 7 1 1 8 9 2 R U (56) Список документов, цитированных в отчете о поиске: RU 2341674 C2, 20.12.2008. US 3301924 A, 27.09.1967. US 4148675 A, 10.04.1979. GB 1084995 A, 27.09.1967. RU 2578659 C1, 27.03.2016. (54) РАКЕТНЫЙ ДВИГАТЕЛЬ ТВЕРДОГО ТОПЛИВА (57) Реферат: Изобретение относится к ракетным промежуточный слой из металлической фольги, двигательным установкам, а именно ...

Solid propellant rocket engine nozzle block insert made of carbon-silica composite material

Изобретение относится к ракетным двигателям твердого топлива (РДТТ), а именно к ракетным соплам, и может быть использовано в сопловом блоке РДТТ пассивного регулирования тяги. Вкладыш соплового блока ракетного двигателя твердого топлива содержит трехмерный объемный каркас, который сплетен из комбинированной нити, состоящей из углеродных и кремнеземных нитей и наполнен пироуглеродом. Изменение соотношения долей углеродных и кремнеземных нитей в объемном каркасе способствует приращению площади критического сечения вкладыша от 50 до 100% начального значения. Изобретение обеспечивает поддерживания постоянного уровня давления продуктов сгорания, получения оптимальных внутри баллистических характеристик и применения ракетных топлив с температурой горения более 4000 K за счет выполнения вкладыша соплового блока из материала с регулируемой и планируемой эрозионной стойкостью. 1 з.п. ф-лы, 1 табл., 1 ил. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 746 081 C1 (51) МПК F02K 9/97 (2006.01) F02K 9/32 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/97 (2021.01); F02K 9/32 (2021.01) (21)(22) Заявка: 2020106630, 11.02.2020 (24) Дата начала отсчета срока действия патента: Дата регистрации: 06.04.2021 (45) Опубликовано: 06.04.2021 Бюл. № 10 (54) Вкладыш соплового блока ракетного двигателя твердого топлива из углерод-кремнеземного композиционного материала (57) Реферат: Изобретение относится к ракетным каркасе способствует приращению площади двигателям твердого топлива (РДТТ), а именно критического сечения вкладыша от 50 до 100% к ракетным соплам, и может быть использовано начального значения. Изобретение обеспечивает в сопловом блоке РДТТ пассивного поддерживания постоянного уровня давления регулирования тяги. Вкладыш соплового блока продуктов сгорания, получения оптимальных ракетного двигателя твердого топлива содержит внутри баллистических характеристик и трехмерный объемный каркас, который сплетен применения ракетных ...

Structural member exposed to action of high thermal loads in operation and method of its manufacture

FIELD: rocket engineering. SUBSTANCE: proposed structural member 1 has shell forming inner space to pass gas and it is formed by at least first part 5 consisting of inner wall 8, outer wall 9 and at least one cooling channel 11 between walls. End 12 of inner wall of first part of member is connected with second part 6 of member. Place 18 of connection is located at a distance from inner of member. Invention contains description of method of manufacture of such member. EFFECT: increased service life of structural members, and provision of economic and effective method of manufacture of components of jet engine. 33 cl, 15 dwg ÐÎÑÑÈÉÑÊÀß ÔÅÄÅÐÀÖÈß RU (19) (11) 2 289 035 (13) C2 (51) ÌÏÊ F02K 9/64 (2006.01) B23K 15/04 (2006.01) ÔÅÄÅÐÀËÜÍÀß ÑËÓÆÁÀ ÏÎ ÈÍÒÅËËÅÊÒÓÀËÜÍÎÉ ÑÎÁÑÒÂÅÍÍÎÑÒÈ, ÏÀÒÅÍÒÀÌ È ÒÎÂÀÐÍÛÌ ÇÍÀÊÀÌ (12) ÎÏÈÑÀÍÈÅ ÈÇÎÁÐÅÒÅÍÈß Ê ÏÀÒÅÍÒÓ (21), (22) Çà âêà: 2004122402/06, 15.11.2002 (72) Àâòîð(û): ÕÅÃÃÀÍÄÅÐ ßí (SE) (24) Äàòà íà÷àëà îòñ÷åòà ñðîêà äåéñòâè ïàòåíòà: 15.11.2002 (73) Ïàòåíòîîáëàäàòåëü(è): ÂÎËÜÂÎ ÀÝÐÎ ÊÎÐÏÎÐÅÉØÍ (SE) R U (30) Êîíâåíöèîííûé ïðèîðèòåò: 18.12.2001 SE 0104273-8 18.12.2001 US 60/340,490 (43) Äàòà ïóáëèêàöèè çà âêè: 10.07.2005 2 2 8 9 0 3 5 (45) Îïóáëèêîâàíî: 10.12.2006 Áþë. ¹ 34 (56) Ñïèñîê äîêóìåíòîâ, öèòèðîâàííûõ â îò÷åòå î ïîèñêå: US 6107596 A, 22.08.2000. RU 2158666 C2, 10.11.2000. RU 2061890 Ñ1, 10.06.1996. DE 3535779 Ñ1, 09.04.1987. (85) Äàòà ïåðåâîäà çà âêè PCT íà íàöèîíàëüíóþ ôàçó: 19.07.2004 2 2 8 9 0 3 5 R U (87) Ïóáëèêàöè PCT: WO 03/052255 (26.06.2003) C 2 C 2 (86) Çà âêà PCT: SE 02/02085 (15.11.2002) Àäðåñ äë ïåðåïèñêè: 101000, Ìîñêâà, Ì.Çëàòîóñòèíñêèé ïåð., 10, êâ.15, "ÅÂÐÎÌÀÐÊÏÀÒ", ïàò.ïîâ. È.À.Âåñåëèöêîé, ðåã. ¹ 11 (54) ÏÎÄÂÅÐÆÅÍÍÛÉ ÂÎ ÂÐÅÌß ÐÀÁÎÒÛ ÂÎÇÄÅÉÑÒÂÈÞ ÂÛÑÎÊÈÕ ÒÅÏËÎÂÛÕ ÍÀÃÐÓÇÎÊ ÝËÅÌÅÍÒ ÊÎÍÑÒÐÓÊÖÈÈ È ÑÏÎÑÎÁ ÅÃÎ ÈÇÃÎÒÎÂËÅÍÈß (57) Ðåôåðàò: Èçîáðåòåíèå îòíîñèòñ ê ðàêåòíîé òåõíèêå. Ýëåìåíò (1) êîíñòðóêöèè èìååò îáîëî÷êó, îáðàçóþùóþ âíóòðåííþþ ïîëîñòü äë ïðîõîäà ãàçà, è îáðàçîâàí ïî ìåíüøåé ìåðå ïåðâîé ÷àñòüþ (5), ñîñòî ùåé èç ...

Propulsion system

FIELD: rocket equipment. SUBSTANCE: invention relates to rocket engineering and can be used in propulsion systems operating on solid fuel and autonomous onboard power sources. Propulsion system comprises chamber with nozzle bosses, in holes of which there are inserts with nozzles, screen, powder charge and igniter. Inserts are made composite – in the form of a holder placed along the axis of each sleeve of the bushing with a nozzle, the length of which is less than the length of the holder. On the bushing on the side of the chamber there is a flange, and the bushing itself is installed with axial and radial gaps relative to the holder and is fixed to it with a heat-shielding material. Heat-shielding material is applied on the inner surface of the holder along its entire length, and the supercritical part of the nozzle of the insert is partially made in the bushing, and partially in the heat-shielding material. Inserts are installed in combustion chamber with thrust in its inner surface and fixed by means of shield from heat-shielding material attached to chamber walls and external surface of inserts. Inner surface of the cartridge behind the point of contact with the chamber has a lowering. EFFECT: higher reliability of propulsion unit while maintaining minimum weight of its design. 8 cl, 3 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 698 780 C1 (51) МПК F02K 9/30 (2006.01) F02K 9/97 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/30 (2019.05); F02K 9/974 (2019.05) (21)(22) Заявка: 2018130060, 20.08.2018 (24) Дата начала отсчета срока действия патента: Дата регистрации: 29.08.2019 (73) Патентообладатель(и): Акционерное общество "Конструкторское бюро приборостроения им. академика А.Г. Шипунова" (RU) (45) Опубликовано: 29.08.2019 Бюл. № 25 2 6 9 8 7 8 0 R U (56) Список документов, цитированных в отчете о поиске: US 4150540 A, 24.04.1979. US 3048970 A, 14.08.1962. RU 2290524 C1, 27.12.2006. RU 2189483 ...

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7 ВУ’” 2016114320” АЗ Дата публикации: 17.07.2019 Форма № 18 ИЗИМ-2011 Федеральная служба по интеллектуальной собственности Федеральное государственное бюджетное учреждение ж 5 «Федеральный институт промышленной собственности» (ФИПС) ОТЧЕТ О ПОИСКЕ 1. . ИДЕНТИФИКАЦИЯ ЗАЯВКИ Регистрационный номер Дата подачи 2016114320/07(022482) 13.04.2016 Приоритет установлен по дате: [ ] подачи заявки [ ] поступления дополнительных материалов от к ранее поданной заявке № [ ] приоритета по первоначальной заявке № из которой данная заявка выделена [ ] подачи первоначальной заявки № из которой данная заявка выделена [ ] подачи ранее поданной заявки № [Х] подачи первой(ых) заявки(ок) в государстве-участнике Парижской конвенции (31) Номер первой(ых) заявки(ок) (32) Дата подачи первой(ых) заявки(ок) (33) Код страны 1. 14/805,259 21.07.2015 05 Название изобретения (полезной модели): [Х] - как заявлено; [ ] - уточненное (см. Примечания) СИСТЕМЫ, УСТРОЙСТВА И СПОСОБЫ ПОДАВЛЕНИЯ ВОСПЛАМЕНЕНИЯ Заявитель: Зе Боинг Компани, 05 2. ЕДИНСТВО ИЗОБРЕТЕНИЯ [Х] соблюдено [ ] не соблюдено. Пояснения: см. Примечания 3. ФОРМУЛА ИЗОБРЕТЕНИЯ: [Х] приняты во внимание все пункты (см. п см. Примечания [ ] приняты во внимание следующие пункты: р [ ] принята во внимание измененная формула изобретения (см. Примечания) 4. КЛАССИФИКАЦИЯ ОБЪЕКТА ИЗОБРЕТЕНИЯ (ПОЛЕЗНОЙ МОДЕЛИ) (Указываются индексы МПК и индикатор текущей версии) Е020 41/22 (2006.01) А62С 2/06 (2006.01) 5. ОБЛАСТЬ ПОИСКА 5.1 Проверенный минимум документации РСТ (указывается индексами МПК) [020 41/22 Аб2С 2/06 5.2 Другая проверенная документация в той мере, в какой она включена в поисковые подборки: 5.3 Электронные базы данных, использованные при поиске (название базы, и если, возможно, поисковые термины): РУ\У/РТ, Ебрасепес, Рабзеагсв 6. ДОКУМЕНТЫ, ОТНОСЯЩИЕСЯ К ПРЕДМЕТУ ПОИСКА Кате- Наименование документа с указанием (где необходимо) частей, Относится к гория* относящихся к предмету поиска пункту формулы № 1 2 3 А 05$ 2015060465 АТИЛМВАСНЕК ВЕКМ№ ...

Solid fuel rocket engine

FIELD: rocket equipment.SUBSTANCE: invention relates to the field of rocket equipment and can be used in rocket engines of solid fuel and autonomous on-board power sources. Solution of the set task is achieved by the fact that in solid-propellant rocket engine consisting of combustion chamber, partially armoured powder charge and nozzle bottom, in the centre of which there is a cavity communicating with the charge end, in the wall separating the cavity and the volume between the charge and the bottom with the nozzles, there are evenly distributed gas ducts on its surface connecting the cavity with the volume between the charge and the nozzle bottom. Some gas ducts are made in plane of nozzles. Total flow area of the channels is determined by the ratio, where F- total flow area gas-water channels m; Sis the powder charge combustion surface area limited by the cavity dimensions, m; Fis area of critical section of nozzles of solid propellant rocket engine, m; Sis area of initial surface of powder charge burning, m.EFFECT: higher reliability of functioning of rocket engine of solid fuel with propellant charge, having developed initial burning surface.1 cl, 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 727 116 C1 (51) МПК F02K 9/34 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/34 (2020.02) (21)(22) Заявка: 2019126445, 22.08.2019 (24) Дата начала отсчета срока действия патента: Дата регистрации: 20.07.2020 (45) Опубликовано: 20.07.2020 Бюл. № 20 (56) Список документов, цитированных в отчете о поиске: RU 2267024 C1, 27.12.2005. RU 2527903 C1, 10.09.2014. US 4150540 A, 24.04.1979. US 3180086 A, 27.04.1965. US 3253407 A, 31.05.1966. 2 7 2 7 1 1 6 R U (54) Ракетный двигатель твердого топлива (57) Реферат: Изобретение относится к области ракетной техники и может найти применение в ракетных двигателях твердого топлива и автономных бортовых источниках энергии. Решение поставленной задачи достигается тем, что в ...

Improved thermal insulation of rocket engines

본 발명은 로켓 엔진에 사용을 위한 배기관(6)의 개선된 단열재에 관한 것이다. 단열재는 층상 구조로 제공되는데, 층들 각각은 합성재를 포함하고, 인접층들 각각의 섬유들의 주 방향들은 서로에 대하여 다르다. 덧붙여서, 유연한 층이 배기관(6)의 내벽에 단열재를 첨부하기 위해 사용될 수 있다.

REACTIVE AMMUNITION ENGINE

РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2019 126 088 A (51) МПК F02K 9/95 (2006.01) F02K 9/32 (2006.01) F02K 9/36 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2019126088, 16.08.2019 (71) Заявитель(и): Российская Федерация, от имени которой выступает Министерство обороны Российской Федерации (RU) Приоритет(ы): (22) Дата подачи заявки: 16.08.2019 (43) Дата публикации заявки: 16.02.2021 Бюл. № 5 A 2 0 1 9 1 2 6 0 8 8 R U Стр.: 1 A (57) Формула изобретения 1. Двигатель реактивного боеприпаса, содержащий камеру сгорания с размещенным в ней трубчатым пороховым зарядом щеточной конструкции, отделенным от воспламенителя диафрагмой, имеющей сквозные отверстия, и сопло, отличающийся тем, что разделительная диафрагма в камере установлена подвижно, а на ее внешнем контуре имеются равномерно расположенные отверстия, образующие по периметру диафрагмы выступы в виде зубьев, при этом отношение суммарной площади отверстий к площади описанного по зубьям круга составляет от 0,279 до 0,339, а отношение внешнего диаметра диафрагмы к внутреннему диаметру камеры сгорания находится в пределах 0,997 - 1,028. 2. Двигатель реактивного боеприпаса по п. 1, отличающийся тем, что разделительная диафрагма выполнена из неметаллического материала. 3. Двигатель реактивного боеприпаса по п. 2, в котором неметаллический материал, из которого выполнена разделительная диафрагма, представляет собой картон толщиной не менее 1 мм. 2 0 1 9 1 2 6 0 8 8 (54) ДВИГАТЕЛЬ РЕАКТИВНОГО БОЕПРИПАСА R U Адрес для переписки: 119160, Москва, Фрунзенская наб., 22/2, Управление интеллектуальной собственности, военно-технического сотрудничества и экспертизы поставок вооружения и военной техники Министерства обороны Российской Федерации (72) Автор(ы): Руссков Владимир Федорович (RU), Середа Николай Владимирович (RU), Аспидов Рудольф Иванович (RU), Степин Валерий Валентинович (RU)

Pulsed solid-fuel engine

FIELD: rocket technology. SUBSTANCE: invention relates to rocket engineering and can be used in making solid fuel pulsed engines, to which raised requirements of different impulses during operation in pair or in whole cluster are applied. Pulsed solid-propellant engine comprises combustion chamber with charge from all-round burning cylindrical channel cartridges, located between support grates, nozzle, igniter, fixed on front support grate on combustion chamber bottom part side, and explosive cartridge installed in combustion chamber bottom part. Between nozzle and support grate located on nozzle side, elliptic shape perforated thin-walled heat-resistant partition is installed, with convex surface facing towards nozzle and with perforation in form of through holes. Partition through holes axis make acute angle with nozzle axis with top towards nozzle throat. Partition holes total area exceeds nozzle throat area. EFFECT: invention reduces spread of solid-fuel pulsed engine burst of power due to increased duration of fuel particles residing in nozzle path. 3 cl, 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 604 772 C1 (51) МПК F02K 9/34 (2006.01) F02K 9/36 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2015127383/06, 08.07.2015 (24) Дата начала отсчета срока действия патента: 08.07.2015 (45) Опубликовано: 10.12.2016 Бюл. № 34 2 6 0 4 7 7 2 R U (54) ТВЕРДОТОПЛИВНЫЙ ИМПУЛЬСНЫЙ ДВИГАТЕЛЬ (57) Реферат: Изобретение относится к области ракетной расположенной со стороны сопла, установлена техники и может быть использовано при создании перфорированная тонкостенная термостойкая твердотопливных импульсных двигателей, к перегородка эллиптической формы, обращенная которым предъявляются повышенные требования выпуклой поверхностью к соплу и имеющая разноимпульсности при работе в паре или в целой перфорацию в виде сквозных отверстий. Оси связке. Твердотопливный импульсный двигатель сквозных отверстий перегородки ...

High-precision solid propellant gas generation amount testing device and method

一种高精度固体推进剂燃气生成量测试装置及方法，属于推进剂理化特性分析测试领域；所述装置包括燃烧室、加热装置、压强传感器、真空泵、气瓶和测试计算机；燃烧室的出口端通过三通接口分为两路，一路依次接压强传感器和测试计算机，用于实时监测所述燃烧室内压强数据；另一路通过高压针阀分别与真空泵、气瓶连接，分别用于所述燃烧室的抽真空和充氮气；燃烧室的燃烧室壳体、燃烧室端盖和泄气罩之间均采用螺纹密封连接；加热装置包括恒温油浴槽、搅拌磁子、加热介质、温度传感器和底座，燃烧室由恒温油浴槽逐渐加热，从而使推进剂样品受热自燃，排除了点火丝对实验测试结果的干扰，同时有利于实验装置的简易化，提升了实验的操作性和可靠性。

Solid propellant rocket engine

Ракетный двигатель твердого топлива, содержащий корпус, во внутренней полости которого размещен заряд, сопло, переднюю крышку, выполненную в виде стакана, с внутренней цилиндрической поверхностью которого контактирует поршень, установленный с возможностью продольного перемещения, на поршне посредством узлов фиксации закреплен полезный груз, причем между поршнем и дном стакана установлен аккумулятор давления, а на открытом торце стакана установлен упорный буртик, отличающийся тем, что аккумулятор давления рассчитан на создание давления в стакане, превышающего давление, на которое рассчитана обечайка стакана, на величину, равную или меньшую давления, создаваемого во внутренней полости корпуса за счет горения заряда. РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (51) МПК F02K 9/08 (13) 2015 137 371 A (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ЗАЯВКА НА ИЗОБРЕТЕНИЕ (21)(22) Заявка: 2015137371, 01.09.2015 (71) Заявитель(и): Зиньковский Александр Тихонович (RU) Приоритет(ы): (22) Дата подачи заявки: 01.09.2015 (43) Дата публикации заявки: 10.02.2016 Бюл. № 04 (72) Автор(ы): Зиньковский Александр Тихонович (RU) A R U A 2 0 1 5 1 3 7 3 7 1 (57) Формула изобретения Ракетный двигатель твердого топлива, содержащий корпус, во внутренней полости которого размещен заряд, сопло, переднюю крышку, выполненную в виде стакана, с внутренней цилиндрической поверхностью которого контактирует поршень, установленный с возможностью продольного перемещения, на поршне посредством узлов фиксации закреплен полезный груз, причем между поршнем и дном стакана установлен аккумулятор давления, а на открытом торце стакана установлен упорный буртик, отличающийся тем, что аккумулятор давления рассчитан на создание давления в стакане, превышающего давление, на которое рассчитана обечайка стакана, на величину, равную или меньшую давления, создаваемого во внутренней полости корпуса за счет горения заряда. 2 0 1 5 1 3 7 3 7 1 (54) Ракетный двигатель твёрдого топлива R U Адрес для переписки: ...

A kind of flow for high-strength hydrogen peroxide change propulsive solid-liquid rocket positions adjustable DC ejector filler

本发明公开一种适用于高浓度过氧化氢变推力固液火箭发动机的流量定位可调直流喷注器，包括头盖、喷注器壳体、阀芯、上盖、入口接头。所述头盖顶部安装入口接头；固定于头盖内腔中，与头盖间具有间隙；喷注器壳体外壁上设有推进剂入口，底端开有推进剂出口；喷注器壳体顶部还安装有上盖。阀芯设置于喷注器壳体内腔中；由入口接头进入头盖内腔中的推进剂，经推进剂入口进入喷注器壳体内腔，通过弹簧和气液压差驱动阀芯动作，阀芯与推进剂出口分离，推进剂喷出。本发明简单可靠、喷注压降小、尺寸小、响应快；通过与可调文氏管配合，可以在流量调节比较大时保持较小的喷注压降的变化。

Rocket engine with solid fuel

FIELD: machine building; cosmonautics. SUBSTANCE: invention relates to solid-propellant rocket engines (SPRE). In a solid-propellant rocket engine comprising a housing from a composite material, comprising a bottom with a metal flange disposed in the central opening of the bottom, and connected to metal flange nozzle with gas duct, support flange is installed on flange with support on flange annular ledge surface, in the annular groove of which on the inner surface on the side of the outer surface of the housing bottom a clamping ring is installed in an axially movable manner, which is borne against the external surface of the housing by the rubber gasket, wherein that in the support ring there are a number of through threaded holes arranged along coaxial support ring of ring, in which there are bolts, wherein bolts are screwed against the stop in the clamping ring. EFFECT: proposed technical solution makes it possible to improve the SPRE structure operational reliability with the nozzle having the gas duct. 6 cl, 3 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 723 276 C1 (51) МПК F02K 9/34 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/34 (2020.02) (21)(22) Заявка: 2019116425, 28.05.2019 (24) Дата начала отсчета срока действия патента: Дата регистрации: 09.06.2020 (45) Опубликовано: 09.06.2020 Бюл. № 16 2 7 2 3 2 7 6 R U (54) Ракетный двигатель на твёрдом топливе (57) Реферат: Изобретение относится к ракетным двигателям твердого топлива (РДТТ). В ракетном двигателе на твердом топливе, содержащем корпус из композиционного материала, включающий днище с металлическим фланцем, расположенным в центральном отверстии днища, и соединенное с металлическим фланцем сопло с газоходом, на фланец с опорой на поверхность кольцевого уступа фланца установлено опорное кольцо, в кольцевой проточке которого на внутренней поверхности со стороны наружной поверхности днища корпуса установлено Стр.: 1 (56) Список ...

Materials and products capable of resisting high temperatures in oxidising medium and method for manufacturing thereof

FIELD: chemistry. SUBSTANCE: invention relates to obtaining material, capable of resisting high temperatures in oxidising medium, and can be used in manufacturing construction parts and coatings. Fireproof material contains at least first component, corresponding to hafnium or non-oxide hafnium compound, or their mixture; second component, corresponding to boron, or non-oxide boron compound, or corresponding mixture of boron and non-oxide boron compound; and third component, corresponding to rare earth element RE or non-oxide compound of rare earth element RE, or corresponding mixture of rare earth element RE and non-oxide compound of rare earth element RE, with RE being selected from scandium, yttrium and lanthanides. Said material is capable of forming self-reducing liquid phase in form of rare earth metal oxide at temperature of exploitation higher than 2000°C. Material contains neither silicon nor silicon compound. EFFECT: preservation of high mechanical properties of parts from fireproof material at temperatures higher than 2000°C and provision of effective protection of parts from oxidation. 9 cl, 9 tbl, 7 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 579 054 C2 (51) МПК C04B 35/58 (2006.01) C04B 41/87 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ 2013130211/03, 05.12.2011 (24) Дата начала отсчета срока действия патента: 05.12.2011 Приоритет(ы): (30) Конвенционный приоритет: (43) Дата публикации заявки: 20.01.2015 Бюл. № 2 (56) Список документов, цитированных в отчете о поиске: C.-M.CHEN et al, "High Temperature Oxidation of LaB6-ZrB2 Eutectic in situ Composite", Acta Material, 1999, vol 47, N 6, p.1945-1952. COURTRIGHT E.L. et al "Oxidation of hafnium carbide and hafnium carbide with additions of tantalum and praseodymium", Oxidation of metals, 1991, v.36, N5-6. ГУЗМАН И.Я., ред. "Химическая технология керамики", Москва, (см. прод.) 2 5 7 9 0 5 4 (85) Дата начала рассмотрения заявки PCT на ...

Rocket combustion chamber proviced with porous layer in its outer shell and manufacturing method thereof

(57)【要約】本公報は電子出願前の出願データであるた め要約のデータは記録されません。

Jet engine

FIELD: engine devices and pumps. SUBSTANCE: jet engine comprises a set of thin alternating washers, detonator rings and an igniting rod. The washers are made of two types: the first type of the washers is made of explosive material, and the second type of the washers is made with beads and latches along the edges, isolating, which is the body of the rocket engine. In the center of the holes in the set of washers there is a igniting rod of the Bengal fire type, burning at a certain rate, on which, at the level of all the washers from the explosive material, ring detonators are installed. The rings detonators are made of two components: the inner layer of the material of the igniting rod, and the outer layer of the explosive mixture. EFFECT: invention makes it possible to simplify the design of the jet engine and to eliminate the difficulties when dumping the waste parts, to provide separation of the engine in emergency situations. 3 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 658 155 C2 (51) МПК F02K 9/94 (2006.01) F02K 9/32 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ОПИСАНИЕ ИЗОБРЕТЕНИЯ К ПАТЕНТУ (52) СПК F02K 9/94 (2016.08); F02K 9/32 (2016.08) (21)(22) Заявка: 2014129997, 21.07.2014 (24) Дата начала отсчета срока действия патента: (73) Патентообладатель(и): Саулин Виктор Леонидович (RU) Дата регистрации: 19.06.2018 (56) Список документов, цитированных в отчете о поиске: US 3811380 A, 21.05.1974. US (43) Дата публикации заявки: 10.02.2016 Бюл. № 4 3815359 A, 11.06.1974. US 3701256 A, 31.10.1972. US 3889462 A, 17.06.1975. RU 3789 U1, 16.03.1997. (45) Опубликовано: 19.06.2018 Бюл. № 17 R U (54) ДВИГАТЕЛЬ РЕАКТИВНЫЙ (57) Реферат: Изобретение может использоваться для вывода на орбиту спутников, в реактивных системах залпового огня, для дополнительного разгона артиллерийских снарядов после выхода снаряда из ствола орудия при подлете к цели. Двигатель реактивный состоит из набора тонких чередующихся шайб, колец детонаторов и запального стержня. ...

Solid fuel grain igniter from composite materials

FIELD: machine building. SUBSTANCE: casing of the solid fuel grain igniter from composite materials, comprises a cylindrical shell with an outer heat-insulating coating. The cylindrical shell has a flat bottom on one side and a free end with an internal thread on the other. The free end is closed with dome shaped removable cap with nozzle openings at the top and support-protective grid in the inner cavity. The threaded part of the cap is designed as a sleeve connected to the cap by an adhesive and having a layer fabric reinforcement perpendicular to the thread axis. The cap is entirely made of reinforced fabric with eyelet and a thread, which is an extension of the sleeve thread. EFFECT: invention allows for the improvement of the solid fuel grain igniter reliability. 2 dwg РОССИЙСКАЯ ФЕДЕРАЦИЯ (19) RU (11) (13) 2 611 115 C1 (51) МПК F02K 9/34 (2006.01) F02K 9/95 (2006.01) ФЕДЕРАЛЬНАЯ СЛУЖБА ПО ИНТЕЛЛЕКТУАЛЬНОЙ СОБСТВЕННОСТИ (12) ФОРМУЛА (21)(22) Заявка: ИЗОБРЕТЕНИЯ К ПАТЕНТУ РОССИЙСКОЙ ФЕДЕРАЦИИ 2015142327, 05.10.2015 (24) Дата начала отсчета срока действия патента: 05.10.2015 Дата регистрации: Приоритет(ы): (22) Дата подачи заявки: 05.10.2015 (45) Опубликовано: 21.02.2017 Бюл. № 6 (56) Список документов, цитированных в отчете о поиске: RU 2539939 C1, 27.01.2015. RU 2505696 C1, 27.01.2014. US 4110977 A, 05.09.1978. RU 2443896 C2, 27.02.2012. RU 2127821 C1, 20.03.1999. 2 6 1 1 1 1 5 Адрес для переписки: 141371, Московская обл., Сергиево-Посадский р-н, г. Хотьково, ул. Заводская, ОАО ЦНИИСМ (73) Патентообладатель(и): Открытое акционерное общество Центральный научно-исследовательский институт специального машиностроения (RU) R U 21.02.2017 (72) Автор(ы): Норкин Николай Степанович (RU), Пашутов Аркадий Витальевич (RU), Кульков Александр Алексеевич (RU), Мерзляков Вячеслав Викторович (RU), Базарко Анатолий Николаевич (RU), Концаев Эльвек Валериевич (RU) 2 6 1 1 1 1 5 R U (57) Формула изобретения Корпус воспламенителя заряда твердого топлива из композиционных материалов, ...

Detection system for composite shell contour of large solid rocket engine

本发明公开了一种大型固体火箭发动机的复合壳体轮廓的检测系统，涉及大型固体火箭发动机的复合壳体检测的技术领域，包括：一可绕自身轴线转动的芯模；待检测的壳体，其缠绕在所述芯模上；一可沿平行于所述轴线的方向往复移动的纱架车；检测机构，其包括：‑检测支架，其固设在所述纱架车上；‑至少三个检测探头，其上下布设在所述检测支架上，所有所述检测探头的探测方向均穿过所述轴线；同时，在所述检测探头沿所述轴线的方向移动时，所述检测探头持续检测各自到所述壳体的距离，以根据检测到的所有距离确定所述壳体的轮廓尺寸。本发明检测难度小，速度快，还能够确保检测一致性。

Solid-fuel propulsion device