Abstract: A turbopump system for pumping two separate fluids is disclosed having a r of concentric turbine rotors located within a turbine casing, each directly connected to a pump impeller. The centrifugal pumps, which may be located on opposite sides of the turbine casing are separately connected to their respective fluids such that the two fluids may be pumped simultaneously. A propellant fluid is directed through a generally annular space defined between the concentric turbine rotors to rotate them in opposite directions. Turbine blades attached to each of the turbine rotors extend into this annular space in order to drive the turbine rotors. The pitches of the turbine blades are such that the rotors are driven in opposite directions and one of the turbine rotors may be driven at a different speed than the other.

Abstract: A thruster nozzle sealing system and apparatus is provided for protection of spacecraft thruster motors. The system includes a sealing plug, a sealing plug insertion tool, an outer cover, an outer cover attachment, a ferry flight attachment, and a test stand. The sealing plug provents moisture from entering the thruster engine so as to prevent valve failure. The attachments are interchangeably connectable with the sealing plug. The ferry flight attachment is used during air transportation of the spacecraft, and the outer cover attachment is used during storage and service of the spacecraft. The outer cover provides protection to the thruster nozzle from mechanical damage. The test stand provides a means for testing and certifying the operating condition of each component of the system.

Abstract: A gel/solid bipropellant propulsion system employs fuel-rich combustion gs from a solid gas generator and an oxidizer gel in a highly efficient combustion chamber wherein the fuel-rich combustion gases and the oxidizer gel are each metered through a vortex injector into a combustion chamber to produce a hypergolic reaction. The solid gas generator (SSG) has a preferred composition of glycidly azide polymer (GAP). The GAP SSG is composed of GAP diol and/or triol polymerized with a di-or tri-function isocyanate, such as isophorone diisocyanate. The gel/solid bipropellant propulsion system comprises the SSG in combination with an oxidizer storage/extrusion vessel system, a combustion chamber system, and a system controller which controls initial ignition of the SSG to produce fuel rich combustion gas which pressurize the system. The system controller monitors pressures and flow rates of fuel and gel oxidizer.

Abstract: First turbine stage 14 rotates at low speed, doing moderate work, and driving a low pressure pump (22). Second turbine stage 16 rotates at high speed, doing a major portion of the work, and drives a high pressure pump (28). The low speed pump is compatible with suction requirements, and the second turbine stage does the substantial work at high efficiency.

Abstract: The compact structural assembly comprises a gas generator, and two turbopumps for feeding the combustion chamber of a chamber of a rocket engine with propellants. A main structure which is essentially circularly symmetrical about the axes of rotation of the turbopumps and made of thermostructural composite materials surrounds the turbines and fixes together the body of the gas generator and the pump bodies of the turbopumps. Internal partitioning elements added to the main structure and likewise made of thermo-structural composite materials serve to divide the space inside the main structure into a plurality of intercommunicating cavities allowing hot gases from the gas generator to flow to the turbines, and also serving to collect the outlet gases from the turbines and exhaust them through a common exhaust duct.

Abstract: In the representative embodiment of the new and improved methods and apparatus disclosed herein to provide effective real-time management of a spacecraft rocket engine powered by gaseous propellants, real-time measurements representative of the engine performance are compared with predetermined standards to selectively control the supply of propellants to the engine for optimizing its performance as well as efficiently managing the consumption of propellants. A priority system is provided for achieving effective real-time mangagement of the propulsion system by first regulating the propellants to keep the engine operating at an efficient level and thereafter regulating the consumption ratio of the propellants. A lower priority level is provided to balance the consumption of the propellants so significant quantities of unexpended propellants will not be left over at the end of the scheduled mission of the engine.

Abstract: An arrangement for the supply of propellant for the propulsion systems of a satellite has a main fuel tank, a main oxidizer tank and an additional or auxiliary oxidizer tank. The auxiliary oxidizer tank becomes effective when the oxidizer in the main tank has been used up at the end of the useful life of the satellite. At that time the additional oxidizer is used to bring the satellite into a retirement orbit out of the way, so to speak, to reduce the crowding in the range of desirable orbits. For this purpose, the emptiness of the main oxidizer tank is sensed by a sensor for providing a control signal which closes the main supply line, thereby switching on the auxiliary tank for propelling the satellite into the retirement orbit.

Abstract: A propulsion engine (12) which combusts propellant received from the storage tank (24) in which a portion (28) of the tank contains propellant in the liquid state and in which an ullage (26) in a remaining portion of the tank contains the propellant in a gaseous state including a first propellant circuit (34) coupling liquid propellant stored in the tank to an evaporator (30); a second propellant circuit (36), coupling the gaseous propellant from the evaporator to the propulsion engine combustor and to the ullage; at least one power generating device (14 and 16) disposed in the second propellant circuit between the evaporator and the ullage, for providing a power output from energy of the gaseous propellant flowing in the second propellant circuit controlled by at least one control valve controlling a flow of gaseous propellant to the at least one device under the control of at least one valve control signal; a bypass circuit (40) coupled in parallel with the at least one power generating device containing a

Abstract: A propulsion system (10) having a propulsion engine (18) which combusts propellant received from a storage tank (12) in which a portion (14) of the tank contains propellant in a liquid state and in which an ullage (16) in a remaining portion of the tank contains the propellant in a gaseous state is disclosed. A first propellant circuit couples liquid propellant stored in the portion of the tank storing the propellant in a liquid state to an evaporator (22) for gasifying the liquid propellant. A second propellant circuit couples the gaseous propellant from the evaporator to the propulsion engine for combustion by the engine and to the ullage. A turbine (24) is disposed in the second propellant circuit coupling the gaseous propellant from the evaporator to the propulsion engine and to the ullage between the evaporator and the ullage for providing a power takeoff (26) powered by energy of the gaseous propellant flowing in the second circuit.

Abstract: A thrust engine is disclosed in the form of a hybrid staged combustion-expander topping cycle engine. The engine comprises a mixed cycle, one cycle being an expander cycle (14) operating at low temperatures and the other being a staged combustion cycle (12) operating at a higher temperature. A portion of liquid hydrogen from a fuel pump (16) is passed in heat exchange relation with the engine nozzle (22) for cooling same and heating the hydrogen, and the heated hydrogen is passed to an oxidizer turbine (28) for driving a liquid oxygen pump (81). Another portion of the liquid hydrogen from the fuel pump is passed in heat exchange relation with the engine thrust chamber (21) and combustor (20), and a portion of the resulting heated hydrogen is introduced, together with a portion of the pressurized liquid oxygen from the liquid oxygen pump, into a preburner (20) for combustion therein, and passage of the combustion gasses to a fuel turbine (26) for driving the fuel pump.

Abstract: A redundant rocket engine for high reliability that incorporates a single thrust chamber and its attendent hydrogen and oxygen injectors and provides a pair of independent but indentical engine systems that provide the fuel feed and cooling systems where one system remains in a standby condition until a malfunction occurs while the other system remains operative. The independent systems are comprised of all of the movable parts.

Abstract: Difficulties in starting turbine engines 10 at high altitudes are avoided in a starting system which includes a combustor 34 having a gas outlet 32 adapted to be in communication with a turbine engine nozzle 28, a fuel inlet 36, and an oxidant inlet 38 so that fuel may be oxidized within the combustor 34 by an oxidant to provide hot gases at the outlet 32. The system includes a first pressure vessel 58 adapted to receive liquid oxygen and a second pressure vessel 60 adapted to receive liquid nitrogen. Pumps 62, 64, and conduits 56, 54, 68 establish a flow path between the vessels 58, 60 and the oxidant inlet 38 while a mixing valve 66 is disposed in the flow path to mix the oxygen and nitrogen flowing therein in a desired proportion.

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 spacecraft propulsion system which integrates the function of the apogee kick motor (AKM) and reaction control system (RCS) is disclosed. In accordance with this invention, a pump-fed AKM is employed which results in lightweight main tanks and pressurization systems. The RCS thrusters are operated by small bellows tanks which are intermittently pressurized by a gas pressurization system to provide high pressure for operation of the RCS thrusters. The system according to this invention enables use of lighter weight main propellant tanks since they do not have to withstand high internal pressures and also enables realization of the numerous advantages of a pump-fed AKM. Several embodiments describe various methods for cycling the bellows tanks.

Abstract: The present invention is related to a liquid fuel rocket engine of the type that a liquid fuel is boosted in pressure by way of a booster means.

Abstract: A dual-fuel, dual-mode rocket engine 60 is made by modifying a baseline single-mode booster engine 10. A second hydrogen propellant system 62, and a fuel mixer 68 is added to provide a means for delivering and mixing a hydrogen fuel 66 with the baseline engine hydrocarbon fuel 36 upstream of exhaust nozzle cooling jacket 38. A second dual-fuel, dual-mode rocket engine 61 is made by modifying a baseline single-mode main engine 11. A hydrocarbon propellant system 63, and a fuel mixer 69 is added to provide a means for delivering and mixing a methane fuel 67 with the baseline engine hydrogen fuel 23 upstream of exhaust nozzle cooling jacket 27. The resulting fuel mixture within both embodiments of the invention described above is utilized for thrust chamber fuel and exhaust nozzle cooling. The relative quantities of each fuel within the mixture vary to provide a progressively less dense mixture throughout a rocket flight.

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

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: A rocket propulsion system working with cryogenic fuel is modified by supersonic and hypersonic flight wherein air is sucked in to be used in lieu of the oxidizer in the rocket system. Hybrid operation is provided for.

Abstract: In a liquid-propellant management system for providing substantially gas-free liquid propellant to thrusters of a space vehicle, a storage tank (10) has an opening (11) through which liquid propellant is delivered via a line (13) to the thrusters. A trap (14) is secured within the tank (10) adjacent the opening (11). and an exit port (15) of the trap (14) is aligned with the opening (11) of the tank (10). A liner (16) is secured inside the trap(14) to define a volume between an interior surface portion of the trap (14) and the liner (16), which volume communicates via the exit port (15) and the opening (11) with the line (13). Liquid propellant enters into the trap (14) through an inlet window screen (27), and passes through porous windows (28) in the liner (16) into the volume between the trap (14) and the liner (16). Vanes (19) are secured to the exterior surface of the trap (14) and extend into the interior of the tank (14) and the liner (16).

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: An improved rocket staging system for missiles and the like wherein a carriage borne rocket engine assembly is sequentially employed within separate, generally aligned oxidizer stages which are generally coaxially disposed about the central rocket engine and its associated carriage. A central fuel tank is surrounded by several separate, cooperating, generally ring-shaped oxidizer tanks generally coaxially disposed about the rocket periphery. A plurality of oxidizer delivery lines run through each of the outer tanks and up to the top of the fuel tank, where a flexible hose brings oxidizer down to the engine carriage. As fuel is consumed, the rocket motor carriage slides upwardly inside the fuel tank in response to thrust. When the carriage is firmly seated inside the next higher oxidizer tank and all of the propellant has been removed from the lowest tank, the lowest tank is jettisoned to discard unnecessary mass.

Abstract: A method is provided for controlling the utilization of a fluid bipropellant including two respective constituents separately in respective tanks aboard a spacecraft for consumption by a spacecraft rocket engine, comprising the steps of actuating the rocket engine; during the actuation of the rocket engine, providing a flow of bipropellant constituents to the rocket engine in a first proportion; after the actuation of the rocket engine, measuring the amount of at least one bipropellant constituent in a bipropellant tank containing the constituent; adjusting a pressure level within at least one bipropellant tank relative to a pressure level within another bipropellant tank based upon the amount of said at least one bipropellant constituent in a bipropellant tank containing said at least one constituent; after the measuring step, actuating the rocket engine; during the actuation of the rocket engine after the measuring step, providing a flow of bipropellant constituents to the rocket engine in a second proporti

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.

Abstract: The process comprises the comparison of the flows of fuel and oxidizer measured at the outlet of the turbopumps with a set value corresponding to the optimum value of the mixture. The set value is periodically modified in dependence on the volumes of fuel and oxidizer remaining in the tanks. The control system for carrying out the process comprises a control valve controlled by an electronic device processing the measurements and disposed in a fuel or oxidizer supply pipe of a driving gas generator for the turbine of a turbopump or in the gas supply pipe of a turbopump, or disposed between the two pipes supplying the same fuel or oxidizer to two driving gas generators, or between the gas outlets of the two turbines of the turbopumps.

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: An improvement in the design of a staged-combustion-cycle rocket engine. The preburner 16 (or preburners) is operated at a higher temperature than that above which turbine blades will be damaged. A heat-exchanger unit 40 through which the output flow of the cooling jacket 18 of the engine is passed is placed inside the preburners 16, or in close proximity to its output flow, so that heat energy is transferred from the output flow of the preburners 16 to the output flow of the cooling jacket 18. This lowers the preburners output flow to a temperature which will not damage turbine blades and raises the cooling-jacket output flow temperature. Since the output flow of the cooling jacket 18 is fed to the low-pressure turbines 22, the increased temperature raises the pressure of the low-pressure turbines 22 so that their output of flow rate is increased and the flow rate to the engine is increased, thereby increasing engine output power and efficiency.

Abstract: A heat-regenerative, expander-cycle, turbine-drive rocket engine 26 in which heated oxidizer gas is used to drive the fuel and oxidizer turbines 16 and 20. The hot exhaust gas from the oxidizer turbine 20 is passed through a heat-donor coil 36 of a heat exchanger 24 where it passes heat to the oxidizer flowing through a donee coil 34 to preheat the oxidizer liquid and gasify it before it is passed through the cooling jacket 28 of the rocket engine 26 where it cools the engine 26 and is itself heated to a higher temperature. The oxidizer, e.g., N.sub.2 O.sub.4, is brought to higher temperature and pressure than its supercritical temperature and pressure so that flashing and boiling of the oxidizer are avoided.

Abstract: The process comprises the comparison of the flows of fuel and oxidizer measured at the outlet of the turbopumps with a set value corresponding to the optimum value of the mixture. The set value is periodically modified in dependence on the volumes of fuel and oxidizer remaining in the tanks. The control system for carrying out the process comprises a control valve controlled by an electronic device processing the measurements and disposed in a fuel or oxidizer supply pipe of a driving gas generator for the turbine of a turbopump or in the gas supply pipe of a turbopump, or disposed between the two pipes supplying the same fuel or oxidizer two driving gas generators, or between the gas outlets of the two turbines of the turbopumps.

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: A gas-pressurizer is disclosed. An enclosed canister surrounds a piercing chanism, over the end of a pressure-conducting pipe. The piercing mechanism is comprised of an end-cap with portals therethrough, a stub, and a gas dispensing, piercing pintle riding upon the stub. When pressurized gas is sent through the pipe, it forces the pintle through the upper end of the canister, allowing the gas to pressurize a storage tank.

Abstract: The invention is for a system which provides total cooling for an aircraft airframe which is designed to fly in the speed range of Mach 2 to Mach 8. The system eliminates the necessity of shielding an aircraft airframe constructed of material such as aluminum. Cooling is accomplished by passing a coolant through the aircraft airframe, the coolant acting as a carrier to remove heat from the airframe. The coolant is circulated through a heat pump and a heat exchanger which together extract essentially all of the added heat from the coolant. The heat is transferred to the aircraft fuel system via the heat exchanger and the heat pump. The heat extracted from the coolant is utilized to power the heat pump. The heat pump has associated therewith power turbine mechanism which is also driven by the extracted heat. The power turbines are utilized to drive various aircraft subsystems, the compressor of the heat pump, and provide engine cooling.

Abstract: A liquid rocket propulsion system utilizing vehicle spin forces and/or propellant tank pressure to charge a specific amount of propellant through a differential area piston to create high chamber pressure.

Abstract: A method of operating a liquid fuel rocket engine having at least first and second propellant component pumps, driven by a gas turbine which is connected to discharge its turbine propulsion gases through a secondary thrust nozzle, includes a main engine having a combustion chamber with a thrust nozzle, an initial expansion portion with an expansion corresponding to ground pressure and a final expansion portion connected to the initial expansion portion and terminating in a thrust discharge operating substantially at vacuum and with means for circulating at least one propellant component through cooling channels of the walls of both expansion portions, comprises, directing a first component into the combustion chamber, and directing at least a first portion of a second propellant component into the walls of the initial expansion portion and then into the combustion chamber and directing a second portion of the second propellant component into the walls of the final expansion portion so that they are heated the

Abstract: A composite cycle rocket engine having an inner engine disposed to discharge directly into the nozzle of an outer engine, is described herein.

Abstract: A rocket propellant feed system utilizing a bleed turbopump to supercharge a topping turbopump. The bleed turbopump is of a low pressure type to meet the cavitation requirements imposed by the propellant storage tanks. The topping turbopump is of a high pressure type and develops 60 to 70 percent of the pressure rise in the propellant.

Abstract: Four embodiments of high-thrust rocket engines, each having separate propellant delivery apparatus for each propellant to be delivered to the engine thrust chamber, are illustrated herein. One engine is designed to provide a high thrust using only one fuel and one oxidizer. The other three engine embodiments are all designed to use different propellants having different densities during different portions of a rocket flight. Each of these additional engines includes the basic structure of the first. The first engine is thus not only an effective engine, but also a building block that can be easily modified to provide an engine capable of operating effectively with different propellants.

Abstract: An improved accumulator particularly adapted for use in controlling the pressure of a stream of fluid in its liquid phase utilizing the fluid in its gaseous phase. The accumulator is characterized by a shell defining a pressure chamber having an entry throat for a liquid and adapted to be connected in contiguous relation with a selected conduit having a stream of fluid flowing through the conduit in its liquid phase. A pressure and volume stabilization tube, including an array of pressure-relief perforations is projected into the chamber with the perforations disposed adjacent to the entry throat for accommodating a discharge of the fluid, in either gaseous or liquid phases, while a gas inlet and liquid to gas conversion system is provided for the chamber and connected with a source of the fluid for continuously pressurizing the chamber for controlling the pressure of the stream of liquid.

Abstract: A liquid-fueled rocket has a storage tank for the fuel components of a liquid fuel and the storage tank is able to communicate with the combustion chambers of the rocket via normally closed valves. The storage tank is enclosed and the fuel components only in part fill the latter. A body of gas is confined in the storage tank and has a pressure exceeding that in the combustion chambers. Thus, when the valves are opened, the fuel components are caused to flow from the storage tank to the combustion chambers by virtue of the difference between the pressure of the gas and the pressure in the combustion chambers.

Abstract: An ignition system for rocket engine combustion chambers which are operated by non-hypergolic propellant components and particularly for the pre-combustion chambers of main stream engines of this type, comprises a main combustion chamber having injector means for injecting a propellant component into the combustion chamber for combustion therein. Since the propellant components employed are non-hypergolically reactive, they are ignited by an ignition agent which is produced by the pressure of a first propellant component which is non-hypergolically reactive with the ignition agent into the combustion chamber. The second propellant component which reacts with the ignition agent hypergolically is also introduced into the combustion chamber so that ignition takes place immediately.

Abstract: This invention is a rocket which has a thrust stage located at the lowermost stage and an arbitrary number of propellant stages which are cut off in turn from a propellant stage located at the uppermost stage after exhausting fuel therein.