Abstract: A thermal regenerative injector 10 for a rocket engine that transfers heat from a fuel after turbine operation to the fuel prior to turbine operation, thereby providing a higher chamber pressure and lower fuel pump pressure. The injector 10 has a plurality of elements 44 having concentric channels 78, 80, and 82 separated by sleeves 66, 68, and 70, the fuels flowing therethrough, heat being transferred through a common middle sleeve 68, The cold fuel, after being so heated, is heated further in the typical fashion of routing it through the combustion zone and nozzle cooling jackets 16 and 20.

Abstract: A valve assembly (10) is provided in which a piston assembly (36) functions to control the introduction and expulsion of a fluid from external and internal manifolds (28,30) defined by the valve assembly housing (12) and the piston assembly. Extensions (42,44,46) provided with flexible disc seals (48) depend from a crown portion (40) of the piston assembly. The seals interface with a T-shaped annular partition (62) having a raised lip (68) in the base (66). The raised lip interfaces with the seals to close off fluid exit ports (80) in the base of the valve assembly housing. When functioned, the housing assembly and piston cause separation of the lip and disc allowing the flow of fluid through the valve assembly.

Abstract: A liquid-propellant delivery system using low-pressure propellant storage tanks instead of heavier high-pressure tanks. Liquid propellant material is pumped from the low-pressure tanks by a solenoid-actuated high-pressure pump, and delivered to the rocket engine through a pressure-actuated injector valve.

Abstract: The present invention provides a rocket thrust chamber (30) comprising a combustor (32) with a fixed divergent profile, a nozzle (36) for further expanding the output of the divergent combustor and a high-velocity, throttleable injector (34) for injecting propellants at such an initial velocity and initially rapid, axial burning rate that the injected propellants achieve sonic flow conditions at preselected locations within the divergent combustor according to throttle setting. In operation, the throttle setting fixes the location of the sonic flow line within the divergent combustor, and the ratio of the nozzle exit area with respect to the effective area of the divergent combustor at the sonic location area fixes the expansion ratio for that throttle setting. At launch, the engine delivers a high propellant flow rate and a moderate expansion ratio, while in space, it delivers a low propellant flow rate and a far greater expansion ratio.

Abstract: A low pressure loss injector element A is disclosed for the main combustion chamber of a rocket engine which includes a lox post (40) terminating in a cylindrical barrel (50). Received within the barrel (50) is a lox plug (46) which is threaded in the lox post and includes an interchangeable lox metering sieve (60) which meters the lox into an annular lox passage (56). A second annular gas passage (62) is coaxial with the annular lox passage (56). A cylindrical sleeve (64) surrounds the annular gas passage (62) and includes an interchangeable gas metering sieve (77) having metering orifices through which a hot gas passes into the annular passage (62). The jets which emerge from the annular lox passage and annular gas passage intersect at 80 which is recessed away from the combustion area. Thus, mixing and combustion stability are enhanced.

Abstract: A constant thrust hybrid rocket motor having a valve assembly therein which utilizes a fixed diameter orifice in combination with a valve having a variable sized opening therein in order to provide a regulated flow of oxidizer to the combustion chamber of the rocket motor. By selecting the size of the fixed diameter orifice such that it is slightly less than required to yield oxidizer flow at highest density oxidizer conditions (lowest temperature) and selecting the size of the combination of the fixed orifice and the open position of the variable opening such that it is slightly greater than required at lowest density oxidizer conditions (highest temperature), the operation of the rocket motor can be reliably maintained over a wide temperature range and without experiencing wide pressure variations by regulating the size of the variable opening.

Abstract: A propellant injector for a liquid rocket engine has an injection cylinder and moving piston in the cylinder. The cylinder on one side of the piston opens into the combustion chamber of the engine. The other side of the piston has a coaxial tubular portion projecting therefrom. A pair of valve members have concentric sleeves slidably mounted respectively on the inside and outside surfaces of the tubular portion of the piston. The outer ends of the concentric sleeves and tubular portion extend through an end wall of the injection cylinder. The valves open and close passages extending through the piston by relative axial movement of the sleeves and the piston. Pneumatic actuator means, connected to the outer ends of the sleeves and tubular portion, moves the sleeves axially relative to the piston to open the valves and move the piston toward the end wall of the injection cylinder to force propellant through the passages into the combustion chamber.

Abstract: The invention relates to an ignition system comprising an igniter constituted by a receptacle, an ignition initiator disposed at one end of said receptacle, and a solid charge placed inside the receptacle, the solid charge comprising pyrotechnic compositions and reactive divided metals capable of reacting with one another exothermically, which compositions successively comprise, from the end of the receptacle carrying the initiator, a pyrotechnic ejection composition, a pyrotechnic ignition composition, and reactive divided metals respectively of decreasing activity, wherein the igniter is associated with a combustion chamber containing a fluid constituted by liquid products or formed by mixtures of gases and of liquids or of solid particles, to ensure, by projection of said solid charge within said fluid, the heating or ignition of said fluid.

Abstract: To provide combustion stability in burners over a wide range of flow conditions, a coaxial injector is provided which includes an outer annular sleeve for introduction of one reactant and a hollow pintle with spaced slots for introduction of another reactant to impinge on the first reactant. The slots are staggered in location and size to permit initial interlocking of the reactants and, therefore, increased mixing and improved combustion performance.

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

Abstract: A dual manifold propellant injector having at least four distinct platelets in its configuration. The first two platelets constituting the primary and secondary fuel platelets; the third and fourth platelets constituting the primary and secondary oxidizer platelets. By proper valve throttling, the injector is capable of providing homogeneous, intimately mixed propellants at all thrust levels.

Abstract: Monopropellant thruster has liquid hydrazine distribution nozzles of such structure as to direct streams of liquid hydrazine into hydrazine decomposition catalyst without puddling and without impinging thruster walls. Thruster also has screens to separate the dissociation catalyst into zones to prevent catalyst voids from standing adjacent distributor nozzles.

Abstract: A control system for supplying uncured propellant into the combustion char of a missile. The system includes a moving injector cooperating with a pressurized putty-like uncured propellant. This combination provides the missile with a propulsion system having rapid wide range throttling, shut off and restart capabilities.

Abstract: High-pressure gas generating apparatus includes a gas generator body provided with a combustion chamber. The gas generator body has a primary propellant injector including orifices for injecting a primary fuel and orifices for injecting a primary oxidizing agent, both the orifices being open to the combustion chamber, an igniter for igniting the primary fuel forced out from the injector, and valve means having a secondary fuel injecting orifice for injecting a metal fuel into a high-temperature gas atmosphere to be reacted therewith and containing a secondary oxidizing agent produced by the reaction of the primary fuel injected from the primary propellant injector with the primary oxidizing agent similarly injected.

Abstract: This invention relates to a method and apparatus whereby functioning airborne conventional aircraft and engine combinations may be assisted during flight in attaining above normal airspeeds and/or altitudes; for the purpose of improving their performance during flight. In the method steps, well known apparatus elements, and/or new and useful combinations thereof, are utilized to enable such aircraft to achieve flight advantages, and improved flight performance, not believed to be possible prior to this invention. According to the method, booster apparatus is used initially in acquiring altitude, forward momentum of the aircraft, rotational momentum of its rotor airfoils, and the like. Apparatus used includes oxygen booster systems. Such systems can include forward thrust rockets, rotational boosters, and apparatus for injecting a supplemental oxidizer into a gas turbine engine afterburner.

Abstract: A dual orifice injector assembly for use in a hybrid rocket motor having a oost and sustain mode and capable of controlling liquid flow rate for the two different modes of operation comprising an inlet manifold to distribute the fluid through a sustain injector and boost injector, actuating means for axially moving the sustain injector into a position to close off the boost injector for the sustain mode of operation and a flow deflector to cause the fluid to be dispersed in a fan like pattern at an angle 180.degree. to the axis of the manifold housing.

Abstract: A sustainer-injector in which fuel core and oxidizer ring members have the lots thereof predeterminely aligned by being keyed together through an intermediate pintle member and with the oxidizer ring being adjustably mounted relative to the pintle and fuel slots. The keying is accomplished by means of machine tabs on the pintle member which engage female slots machined on the fuel core and oxidizer ring members. The oxidizer ring is adjustably mounted relative to the fuel core and pintle member by having means at one side thereof for causing the oxidizer ring to be positioned linearly relative to the pintle member and the fuel core member.

Abstract: An electrothermal reactor is provided to control the decomposition of chemical compounds and mixes and generate a gas flow. The reactor embodies a reaction chamber which has an electrical heat source. Means are provided to inject the chemical compounds and mixes into the chamber such that the compounds will vaporize. The heat source raises the temperature within the chamber to the reaction temperature wherein the chemical compounds and mixes react, thereby releasing chemical energy and producing a gas flow of the reaction products.