Abstract: An injection element for a combustion component running on two fuels, includes a first, central injection channel, which is connectable to a first fuel supply for a first fuel, and a second injection channel, which annularly surrounds the first injection channel. The second injection channel is connectable to a second fuel supply for a second fuel. The injection element includes a third injection channel, which annularly surrounds the second injection channel and which is in fluid communication with the first injection channel.

Abstract: A liquid fuel rocket engine with a closed direct auxiliary flow cycle has a combustion chamber connected to each liquid fuel source through a respective main liquid fuel supply line. A turbo pump is provided in each main fuel supply line. The turbo pump is driven by a gas generator which is connected to a first bypass fuel supply line for supplying fuel to the gas generator. An exhaust gas duct (8, 9, 10) leads from the gas generator into at least one main fuel supply line for feeding gas generator exhaust gas into the main fuel supply line upstream of the turbo pump.

Abstract: A centrifugal direct injection engine having an open-ended compartment with a combustion chamber at a closed end thereof and an exhaust nozzle at its open end. The compartment is provided with a narrow portion defining a throat that connects the combustion chamber and the exhaust nozzle. A propellant injection manifold is rotatably mounted within the combustion chamber, the propellant injection manifold being configured for rapid rotation within the combustion chamber. The propellant injection manifold includes an upper disk member, an intermediate disk member and a lower disk member, the disks being joined together to form an integral member with the outer perimeter of the disks being in registration with one another. A fuel is introduced into the combustion chamber through the propellant injection manifold and the nozzles.

Abstract: A rocket engine includes a generally cylindrical annular combustion chamber, an injector attached to an injector end of the combustion chamber, a generally cylindrical annular step collar fitting within the combustion chamber at the injector end thereof, and an attachment of the combustion chamber, the injector, and the step collar. The maximum outer diameter of the step collar is from about 0.020 inches to about 0.024 inches smaller than the inner diameter of the combustion chamber, and has a length of from about 21 to about 31 percent of the length of the combustion chamber. The step collar is made of a material having a melting point greater than that of the combustion chamber. The attachment preferably includes a clip structure joining the combustion chamber and the step collar.

Abstract: A method for using an injector to combust a first fluid and a second fluid comprises a directing the first fluid step and a shrouding step. The directing step comprises directing the first fluid in a central coaxial flow generally perpendicularly through a first fluid passage that extends through an injector faceplate, and into a combustion chamber. In the shrouding step, a length of the central coaxial flow adjacent to the injector faceplate is shrouded with an outer coaxial flow. The shroud is formed by directing a portion of the second fluid through an annular/conical passage in the injector faceplate that surrounds the first fluid passage. The annular/conical passage is arranged such that the outer coaxial flow is directed radially inward and impinges on the central coaxial flow a distance after the central coaxial flow has exited the first fluid passage and after the outer coaxial flow has exited the annular/conical passage.

Abstract: A gas generator in particular for driving liquid-propellant rockets, having, a structural envelope, an outlet pipe, a cover connected to the structural envelope from the side opposite to the outlet pipe. A bush is made on the inner side of the cover. An injector face is mounted on the bush end to form a space between the injector face and the cover. Chambers are made in the injector face and the cover. A spacer, mounted inside the structural envelope, is fastened on the outlet pipe and the bush to form an annular space there between. Mixing modules are fastened in the cover and in the injector face chambers. A propellent feed inlet pipe is fastened to the outer surface of the cover inside which a propellant space is made. An oxidizer feed inlet pipe is fastened inside the structural envelope and communicates with the annular space. The annular space communicates with the space between the cover and the injector face via ports made in the bush.

Abstract: The invention relates to rocket engine manufacturing, and may also be used in aircraft engine manufacturing. A chamber comprises a casing, ignition means and a mixing head. The mixing head consists of an internal injector face, a middle bottom, an external bottom. Bipropellant injectors are fixed in the internal injector face and the middle bottom. A portion of the bipropellant injectors are mounted projecting from the internal injector face, while another portion are recessed in the internal injector face. The ignition means are made from jet injectors mounted in a structural casing behind the internal injector face. Axes of diverging ports of jet injectors are positioned at an acute angle to an outlet from the structural casing and deflected in a circle in a lateral plane from the longitudinal axis of the structural casing in an identical direction. The chamber casing comprises a combustion chamber and a nozzle, made from an external structural envelope and an internal fire wall.

Abstract: The engine comprises a chamber, fuel and oxidizer booster pumps, a turbopump unit, a gas generator, an electrically driven thrust regulator, and a fuel throttle. A programmable starting and ignition system is introduced, which includes a fuel tank, connected to the thrust regulator, and a first ampoule with starting fuel, an output of which through a starting/cut-off valve is connected to fuel injectors of the gas generator. The aforesaid tank is also connected through a second ampoule to the starting fuel and by a jet to autonomous ignition injectors located in the space of the combustion chamber of the engine. An inclined Archimedian screw pump with a gas turbine as an actuator and an oxidizing gas as a working medium is used as an oxidizer booster pump. The gas is discharged after its use in a turbine into an outlet manifold of the booster pump.

Abstract: The invention relates to the field of power machine building and may be used in mixing devices of gas generators for driving gas turbines, in particular, for liquid-propellant rocket engines. A gas generator comprises a cover with a hole, a bottom with a mixing chamber, a housing with an axial duct communicating with the hole in the cover. The housing is fastened in the cover hole and/or in the mixing chamber of the bottom. Grooves are made uniformly along the circumference of the cross section of the housing in the bottom and/or on the housing. The hole in the cover and the axial duct are designed to supply propellant, and the space between the cover and the bottom and the grooves in the bottom and/or on the housing are designed to supply an oxidizer. A propellant space is made in the cover, and the axial duct communicates through the propellant space with the hole in the cover by means of tangential propellant supply holes.

Abstract: A rocket engine comprises first and second rotary injectors for injecting respective fuel and oxidizer propellant components into a first combustion chamber, and the effluent therefrom drives a turbine that rotates the rotary injectors. The mixture within the first combustion chamber is preferably fuel-rich so as to reduce the associated combustion temperature, and the fuel-rich effluent mixes in a second combustion chamber with additional oxidizer injected by a third rotary injector so as to generate a high temperature effluent suitable for propulsion. The rotary injectors are adapted so as to isolate the low pressure propellant supply from the relatively high pressures in the respective combustion chambers.

Abstract: A tripropellant coaxial pintle injector for a rocket engine which enables smooth transitioning between two types of propellants that are adapted to be alternatively mixed with a third type of propellant in a combustion chamber of a rocket engine. The tripropellant coaxial injector includes a pintle for expelling a first propellant into a combustion chamber in a generally radial direction and two or more concentric metering sleeves which define orifices for expelling two or more different propellants into the combustion chamber in an axial direction thus, such a configuration allows for alternative mixing of two propellants with a third propellant in the combustion chamber of a rocket engine.

Abstract: A liquid fuel rocket engine wherein the thrust producing exhaust of the engine is directed out of the engine in an annular fashion providing increased thrust and 360 degree thrust vectoring, while presenting the ability to be adaptable to different fuel/oxidizer mixtures, and prevent the sinusoidal shock wave formation typically associated with bell nozzle type rocket engines.

Abstract: A rocket engine has a combustion chamber formed of an annular wall made of an annular wall material, and an annular insert made of an annular insert material that is of substantially the same thermal expansion coefficient as the annular wall material. The annular insert has an insert outer surface sized to fit within and contact the wall inner surface along less than the wall length. The annular wall and the annular insert are welded together. There is further an attachment between the combustion chamber and an injector. The attachment includes an annular metallic deposit bonded to the combustion chamber, an annular adaptor ring, a weld joint between the adaptor ring and the metallic deposit, and a joint between the adaptor ring and the injector.

Abstract: An injector for liquid fueled rocket engines wherein a generally flat core having a frustoconical dome attached to one side of the core to serve as a manifold for a first liquid, with the core having a generally circular configuration having an axis. The other side of the core has a plurality of concentric annular first slots and a plurality of annular concentric second slots alternating with the first slots, the second slots having a greater depth than said first slots. A bore extends through the core for inletting a second liquid into said core, the bore intersecting the second slots to feed the second liquid into the second slots. The core also has a plurality of first passageways leading from the manifold to the first annular slots for feeding the first liquid into said first slots.

Abstract: A device for fastening an injection element 4 with its foot 5 in a hole of a base plate 1 of an injection head 2 of a rocket engine, the foot 5 is tapered in steps up to its end. A first shoulder 7 of the foot abuts against the base plate 1; a thread (6) screwed into the base plate 1 is present after the first shoulder 7; the thread 6 is followed by a second shoulder 8 and then by a first, smooth cylindrical piece 9. This is followed by a third shoulder 11, on which a solder ring 12 made of a gold alloy is seated. A second, smooth cylindrical piece 13, which has a solder gap 13a between it and the corresponding hole in the base plate 1, follows after the third shoulder 11 up to the end. A process for fastening the injection element 4 provides for the solder ring 12 being first placed on the third shoulder 11, after which the injection element 4 is screwed into the base plate by means of a hexagon 22, and the injection head 2 is subsequently heated to about 1,035.degree. C.

Abstract: An ignition system for a rocket engine which improves the reliability of engine re-ignition by supplying additional oxygen to the ignitor directly from the rocket engines main fuel/oxygen injector elements, thereby eliminating the need for a supplemental oxygen supply system.

Abstract: A hybrid rocket motor having a multitude of radially inward directed miniature hybrid secondary combustion chambers within a single solid fuel block being oxidizer-fed by a single centrally positioned distributor and a system of radially outward-directed feed tubes. The feed tubes are sealed at their distant ends and have multiple orifices opening into the secondary combustion chambers to resemble a multitude of secondary conventional-injection hybrid motors within a single main hybrid motor.

Abstract: An injection head for rocket engines, in the combustion chambers of which one or more liquid propellants are chemically reacted to generate thrust, with a plurality of injection elements as well as with a distribution chamber arranged before these injection elements in terms of flow. The distribution chamber is divided into two chambers connected in series in terms of flow by means of a separator plate arranged at right angles to the longitudinal direction of the engine. The separator plate has a plurality of openings, each with a different flow cross section and with a defined direction of flow, wherein each opening is associated with one or more injection elements.

Abstract: A liquid propellant rocket engine comprising a chamber body connected to a converging/diverging thrust nozzle and to at least one lightly pressurized propellant feed duct, said body including a body bore in which an injector-forming piston can move longitudinally between a rest first position and an operating second position, the piston comprising a piston head and a piston rod, and subdividing the chamber body bore into a combustion chamber in front of the piston, and at least one injection chamber surrounding the piston rod, the chamber body further including a flow section that is closed in said rest position to prevent any propellant being fed from the injection chamber(s) and that is opened when the piston moves from said rest position to said operating position in order to allow the injection chamber(s) to be filled via said flow section directly from the feed duct(s), the propellant(s) being injected into the combustion chamber via injection channels formed between the combustion chamber and the inject

Abstract: A rocket engine burner for throttling over a large thrust range from full thrust to very low thrust. The rocket engine burner comprises a propellant injector assembly, a combustion chamber, and an ignitor. The propellant injector assembly comprises one or more hollow injectors and a secondary propellant manifold in communication with the hollow injectors to provide them with a secondary propellant. The propellant injector assembly also includes a primary propellant manifold and a porous injector that extends across the inlet of the combustion chamber and is in communication with the primary propellant manifold to receive the primary propellant. The primary propellant flows through the porous injector and is injected into the combustion chamber. The porous injector has one or more openings through it which receive the hollow injectors. The secondary propellant flows through the hollow injectors and is injected into the combustion chamber.

Abstract: A liquid propellant rocket engine comprising a chamber body connected to a converging/diverging thrust nozzle and to at least one lightly pressurized propellant feed duct, said body including a body bore in which an injector-forming piston can move longitudinally between a rest first position and an operating second position, the piston comprising a piston head and a piston rod, and subdividing the chamber body bore into a combustion chamber in front of the piston, and at least one injection chamber surrounding the piston rod, the chamber body further including a flow section that is closed in said rest position to prevent any propellant being fed from the injection chamber(s) and that is opened when the piston moves from said rest position to said operating position in order to allow the injection chamber(s) to be filled via said flow section directly from the feed duct(s), the propellant(s) being injected into the combustion chamber via injection channels formed between the combustion chamber and the inject

Abstract: A propulsion plant operating on the basis of a catalytic and/or chemical decomposition of a liquid propellant, is equipped with a so-called aspirator ring (19) that covers the exit openings of the injection pipes (11), thereby screening these openings toward the decomposition chamber. The aspirator ring (19) is further enclosed by an aspirator screen (19A). The aspirator ring is constructed as a vaulted channel (20) in which the discharge bores (21) are arranged in a pattern that assures a uniform distribution of the propellant into the catalyst bed in the decomposition chamber. The aspirator screen (19A) is also formed with a channel and both the channel in the screen (19A) and the channel in the aspirator ring (19) preferably have a semicircular cross-section. The catalyst bed is divided into two separate beds so that the respective housing can also be divided so that subassemblies of the thruster may be built separately and then assembled.

Abstract: Injection element of coaxial design for rocket combustion engines and for operation with two propellants, at least one of which disintegrates into drops with difficulty only, comprising a central body for the more readily disintegrating propellant (oxidant), whose flow channel generates a conical drop distribution at the outlet; a sleeve concentrically surrounding the central body to form an annular flow channel for the less readily disintegrating propellant (fuel); and a flow splitter which splits this propellant flow into a plurality of individual flows. The flow splitter has an even number of passage channels, which are uniformly distributed around the central body and cooperate in pairs in terms of an impact atomization ("impinging"), wherein the axes of two adjacent passage channels each intersect downstream of the flow splitter and upstream of their points of impact on the propellant cone discharged from the central body.

Abstract: Thrust-generating device with rocket engines arranged in pairs opposite each other for multiple ignition as well as individual and group activation and with an amount of fuel sufficient for the duration of the mission. The thrust-generating device is designed as a closed, storable, modular unit. The rocket engines are operated with liquid fuel and liquid oxidizing agent. At least one tank is provided for each fuel component. The tanks are in the filled, pressureless state before the activation of the thrust-generating device. The rocket engines and a fuel distributor of block design are integrated within a central bulkhead. The thrust-generating device is surrounded by a sealed outer jacket and is divided by the bulkhead into two separate chambers. The outer jacket is lined on the inside with a "self-sealing" film. The oxidizing agent is stored in one chamber, and the fuel in the other chamber. The empty spaces in the chambers contain an adsorbent filling.

Abstract: The present invention relates to an injector of tricoaxial structure comprising a body that is circularly symmetrical about an axis of symmetry, said body being provided with a plurality of injection elements having axes parallel to said axis of symmetry and passing through an injection wall fixed to said body and separating said injector from a combustion member to which it is coupled. According to the invention, each of said injection elements comprises a central injection channel fed with a second propellant and an annular injection slot surrounding said central channel and fed with a first propellant, an additional annular injection channel for the second propellant being formed by a space that exists around each injection element between its side wall and said injection wall, such that the sheet of first propellant flowing from said annular slot is engaged between the central jet and the outer sheet, each of which delivers the second propellant.

Abstract: Disclosed is a low flow, low Reynolds number cavitating venturi. This cavitating venturi includes an inlet for receiving a liquid at an upstream pressure and an outlet for discharging the liquid received by the inlet at a downstream pressure. The liquid passes through a converging portion having a converging sidewall which extends from said inlet, through a throat portion having a throat sidewall and a diverging diffuser portion having a diverging sidewall. The cavitating venturi provides a substantially stable liquid flow rate independent of the downstream pressure up to a downstream pressure at least as high as 80% of the upstream pressure at a Reynolds number of 60,000 or less.

Abstract: Current state-of-the art injection methodologies for bipropellant engines ly on either impingement or turbulence to mix the fuel and oxidizer streams. The multiple impinging stream vortex injector combines both mixing schemes into a single injector. Both first stage mixing at point of impingement and second stage mixing or turbulent vortex mixing is accomplished by impinging momentum balanced, tangentially injected propellant streams onto one another. The impingement angles are calculated to yield a resultant stream vector that consists of only a tangential velocity component. The two stages of mixing results in increased engine performance and combustion efficiency.

Abstract: The present invention relates to a parallel-sheet injector provided with a body about an axis to which there is coaxially connected at least one annular injection element comprising a central annular injection slit surrounded by two other annular injection slits. The central annular slit is fed with a second propellant which enters at the top of the injection element level with said central slit and the other annular slits are fed with a first propellant which enters at the base of the injection element, on either side of the central slit, and passes through an annular homogenization cavity from which it flows towards the annular slits surrounding the central injection slit. The invention also relates to injection elements which may be used within the scope of this injector.

Abstract: A method and apparatus for increasing the operating life of a turbopump turbine by reducing the temperature of the combustion products impinging the turbine blades adjacent the roots thereof while increasing the temperature of the combustion products impinging the other portions of the turbine blades to maintain the overall temperature of the combustion products.

Abstract: Injection element of coaxial design for rocket combustion chambers with a combustion chamber pressure of 3 to 30 bar and for operation with two hypergolically reacting propellants, with a central body for the oxidant, whose flow channel has a twisting insert, a cross section contraction and a sharp opening edge, and with a sleeve concentrically surrounding the central body to limit an annular flow channel for the propellant, wherein the flow channel has a circular cylindrical outlet with a sharp opening edge. The flow channel of the central body has a circular cylindrical outlet. The opening edges of the central body and of the sleeve are located approximately in the plane of the front surface of the combustion chamber. A flow divider, which divides the propellant flow into five to twenty equal individual flows is arranged in the flow channel for the propellant. The velocities of flow of the oxidant and of the propellant are between 5 m/sec and 20 m/sec at the point of entry into the combustion chamber.

Abstract: A high speed control apparatus and method utilized in a system having a high pressure fluid supply for producing precisely controlled, amplitude and duration variable thrust are disclosed, the apparatus including a piezoelectric stack responsive to a control signal input, a displacement amplifier operatively associated with the stack, and a high pressure axial valve assembly coupled with the amplifier. A microprocessor control system and energy storage and power amplifier system provide the calibrated output voltage control signal which drives the piezoelectric stack in such a manner as to produce rapidly changing displacements in the stack (as frequent as 10 microseconds) which are subsequently amplified by the displacement amplifier. The valve assembly opens as the piezoelectric stack expands upon application of a step change in voltage, allowing fluid to expand through the valve assembly and create thrust that is substantially equal in duration and directly proportional in amplitude to the control signal.

Abstract: The rocket engine is fitted with an injection device comprising a single injection plate having a central portion that defines the end of the burn zone of the combustion chamber and that contains first propellant injectors having a high mixing ratio adapted to the conditions under which hot gases are produced in the burn zone, and a peripheral portion in the form of a ring that defines the end of an annular cavity of a gas generator and that contain second propellant injectors having a low mixing ratio and adapted to the conditions under which hot gases are produced in the gas generator. The gas generator is thus integral with the combustion chamber and the annular cavity of the gas generator is separated from the burn zone of the combustion chamber by a cylindrical wall that extends axially perpendicularly to the injection plate and that is fixed thereto in sealed manner.

Abstract: A high speed control apparatus and method utilized in a system having a high pressure fluid supply for producing precisely controlled, amplitude and duration variable thrust are disclosed, the apparatus including a co-fired piezoelectric stack connected with a control signal input, a low-loss mechanical displacement amplifier operatively associated with the stack, and a high pressure spring-loaded axial valve assembly coupled with the amplifier. A microprocessor control system and energy storage and power amplifier system are utilized to provide a calibrated output voltage control signal which drives the piezoelectric stack in such a manner as to produce rapidly changing displacements in the stack (as frequent as 10 microseconds) which are subsequently amplified by the displacement amplifier. The valve assembly is connected with the amplifier and includes a valve core which is normally forced closed upon a valve seat by a compression spring.

Abstract: Thrust chambers for liquid propellant rocket engines include three principal components. One of these components is an injector which contains a plurality of injection elements to meter the flow of propellants at a predetermined rate, and fuel to oxidizer mixture ratio, to introduce the mixture into the combustion chamber, and to cause them to be atomized within the combustion chamber so that even combustion takes place. Evolving from these injectors are tube injectors. These tube injectors have injection elements for injecting the oxidizer into the combustion chamber. The oxidizer and fuel must be metered at predetermined rates and mixture ratios in order to mix them within the combustion chamber so that combustion takes place smoothly and completely. Hence tube injectors are subject to improvement.

Abstract: A device for tapping off hot gases in a combustion chamber into which arrive a plurality of injectors has a tap-off means intended for conveying gases tapped off from the combustion chamber to an additional cavity. The tap-off means is formed of at least one tubular element having, in the region of its end opposite the combustion chamber a nozzle for individual adjustment of the flow of gases tapped off. A number of tap-off devices can be incorporated in an injector head comprising several injectors. The tap-off means comprises a cooling system using one of the cold propellants.

Abstract: A combustor element face plate made of a thermally highly conductive material, such as copper, has a plurality of cooling holes extending therethrough. Some of the cooling holes are located about the injector openings, while other cooling holes are located adjacent the peripheral edge of the face plate.

Abstract: There is disclosed in combination with a solid fuel-propellant grain 24 having an axially extending aperture 26, particularly of a hybrid rocket motor, an oxidizer inlet 20 at one end and a combustion gas outlet 22 at the other end, a tubular oxidizer injector 32 disposed in the inlet and extending axially in the grain aperture. The tubular injector comprises a tube 34 containing a plurality of oxidizer injection orifices 38 in the outer circumference of the tube to discharge fluid streams of oxidizer into the fuel grain aperture for combustion of the fuel grain. Additional oxidizer orifices, e.g. in the form of shower head orifices 44, are disposed in the downstream end of the tube.

Abstract: A rocket injector is provided with multiple sets of manifolds for supplying propellants to injector elements. Sensors transmit the temperatures of the propellants to a suitable controller which is operably connected to valves between these manifolds and propellant storage tanks. When cryogenic propellant temperatures are sensed only a portion of the valves are opened to furnish propellants to some of the manifolds. When lower temperatures are sensed additional valves are opened to furnish propellants to more of the manifolds.

Abstract: A method for operating a rocket engine by injecting fuel and oxidizer into an elongated combustion chamber in two flows, a core flow where the fuel and oxidizer are intimately mixed and immediately combusted and a peripheral curtain flow which surrounds the core flow and which is in contact with the combustion chamber wall to cool it and limit the heat transfer from the wall to the injector to prevent vapor locks in the injector. To prevent decomposed or partially combusted propellant products from chemically reacting with the chamber wall no mixing of fuel and oxidizer takes place in the curtain flow. The curtain flow is deflected radially inward into the core flow, before decomposed or partially combusted products can come into contact with the wall, into the core flow to fully combust the curtain flow out of contact with the wall. The rocket engine is defined by serially arranged first and second combustion chambers and an injector constructed to form the core and curtain flows.

Abstract: A bi-propellant self-contained propulsion system is provided for powering rockets. A plurality of turbopump assemblies are provided to take liquid propellants from low pressure storage tanks to a substantially higher pressure thrust chamber. Substantially all of the liquid propellants are pressurized and gasified within the plurality of turbopumps. Substantially all of the gasified propellants are then used to drive the turbopumps that pressurize the liquid propellants. Gasification preferably occurs within a preburner internal to the turbopump assembly that combines a small portion of one of the propellants with a substantial portion of the other. The proportions are selected so that gasification of all of the propellants is ensured yet relatively low preburner temperatures are maintained. A multi-stage pintle assembly may be provided to vary the exit-to-throat area ratio of the nozzle. The total thrust and the mixture ratio may be controlled by shutting down some of the turbopumps.

Abstract: A bottom unit for the combustion chamber of a liquid-fueled rocket including a bottom plate held on a swivel bearing and having circumferentially distributed openings for the injection of the two fuel components as well as two separate groups of supply passages which are connected to these fuel injection openings. The swivel bearing consists of a rigid thrust plate and a rigid hanger plate which are held and joined together by an internally situated, first annular body of rubber elastic material and by a second annular body of rubber elastic material surrounding the first at a radial distance therefrom. The one group of supply passages is formed by the cavity surrounded by tthe first annular body and the other group of supply passages by the annular space between the first and the second annular body.

Abstract: A construction for the delivery of propellant constituents to the main combustion chamber of a hypergolic liquid bipropellant rocket engine having a staged combustion cycle incorporating at least one precombustor comprises: a mixer for mixing together prior to delivery to the main combustion chamber a first propellant constituent comprising oxidant together with exhaust gas from the precombustor injector means for injecting the mixture of the oxidant and exhaust gas provided by the mixer into the main combustion chamber in such a manner that the injected mixture forms a recirculation zone inside the main combustion chamber, an inlet into the main combustion chamber for a second propellant constituent comprising fuel, and a delivery channel to the said inlet, the said inlet and delivery channel being disposed laterally of the injector means relative to the main direction of flow through the injector means in such a position that fuel delivered into the main combustion chamber thereby is delivered into the said

Abstract: A method for operating a rocket engine by injecting fuel and oxidizer into an elongated combustion chamber in two flows, a core flow where the fuel and oxidizer are intimately mixed and immediately combusted and a peripheral curtain flow which surrounds the core flow and which is in contact with the combustion chamber wall to cool it and limit the heat transfer from the wall to the injector to prevent vapor locks in the injector. To prevent decomposed or partially combusted propellant products from chemically reacting with the chamber wall no mixing of fuel and oxidizer takes place in the curtain flow. The curtain flow is deflected radially inward into the core flow, before decomposed or partially combusted products can come into contact with the wall, into the core flow to fully combust the curtain flow out of contact with the wall. The rocket engine is defined by serially arranged first and second combustion chambers and an injector constructed to from the core and curtain flows.

Abstract: A fluid-powered actuator has a piston slidably mounted in a cylinder. The piston subdivides the cylinder into first and second end chambers. A plurality of solenoid valves are arranged to selectively inject fuel and oxidizer into either chamber, or to inject liquid oxidizer to react with a solid fuel. Upon contact with one another, the fuel and oxidizer undergo a hypergolic reaction, and the products of this reaction create a pressure sufficient to displace the piston. Alternatively, a mono-propellant fluid fuel may be injected into such chamber. Upon contact with a catalyst in the chamber, the fuel undergoes a disassociation reaction, and the products thereof may be used to displace the piston.

Abstract: An injector assembly 10 for injecting fuel and oxidizer in a combustion chamber 38 of a fluid-fueled rocket engine. Injector assembly 10 comprises a stationary housing 12, a plate member 14 attached to the housing 12, at least one injector subassembly 18 and actuation means 16. Each injector subassembly 18 includes a plug 44 attached to the housing 12 and a movable elongated throttle means 64 connected to the actuation means 16. When the throttle is at least partially open, fuel flows through inlet 20, orifice 110, inner passageway 66, orifice 112 and finally to the combustion chamber 38. Oxidizer flows through inlet 22, orifice 114, outer passageway 68, orifice 116 and finally to the combustion chamber 38. Controlling the flow of propellants at upstream orifices 110,114 controls the propellant pressure at the downstream orifices 112,116 during thrust level changes and corresponding combustion chamber pressure changes of the engine, thereby maintaining a constant stream momentum and combustion efficiency.

Abstract: Fuel for a rocket engine consisting of hydrogen, oxygen and a hydrocarbon is admitted into the combustion zone through an injector whereby the hydrogen and hydrocarbon components are either pre-mixed in the injector or independently delivered and the oxygen is always independently delivered into the combustion zone.

Abstract: The invention combines liquid hydrogen with liquid oxygen and a hydrocarbon as the fuels in a tripropellant rocket booster engine.

Abstract: A compact, valve assembly (10) is provided having multiple chambers. A forward chamber (14) retains dual T-shaped blades (36) having shaped orifices (38) for controllably metering at least two fluids passing through the valve assembly. The valve assembly is also provided with a servo-valve assembly (70) and a piston indicator (62) assembly in cooperative association with the valve assembly.

Abstract: A compact pulse-rocket engine assembly comprising an elongated substantially cylindrical housing containing an axially extending substantially cylindrical combustion chamber which terminates at one end in a nozzle means. Within the housing at the opposite end of the combustion chamber are a pair of charging chambers. Intermediate the charging chambers and the combustion chambers are a pair of valve means which are parallel and linked together for ensuring that the valves operate in unison.

Abstract: A compact pulse-rocket engine assembly comprising an elongated housing containing an axially extending substantially cylindrical combustion chamber terminating in a nozzle means at one end. An opposite end of the housing contains a pair of angularly disposed charging chambers. Intermediate the charging chambers and the combustion chamber are a pair of angularly disposed piston valve means. The arrangement of the valve means provides for the injection of propellants into the combustion chamber for impingement at the same point throughout the operation of the piston valve means.