Abstract: The present invention relates to a reactor for the decomposition of ammonium dinitramide-based liquid monopropellants into hot, combustible gases for combustion in a combustion chamber, and more particularly a rocket engine or thruster comprising such reactor and a combustion chamber. The invention also relates to a process for the decompostion of ammonium dinitramide-based liquid monopropellants.

Abstract: A method for forming a coolant system for a rocket engine combustion chamber is provided. The method comprises the steps of providing a plurality of tubes formed and shaped into the profile of a nozzle with each of the tubes having a constantly expanding cross section in an upper chamber area, providing an inlet manifold and an exit manifold with a plurality of holes for receiving an end of each tube, inserting a brazing preform into each hole, inserting a first end of each tube into the inlet manifold and a second end of each tube into the outlet manifold so that the first end is surrounded by a first brazing preform and the second end is surrounded by a second brazing preform, and brazing the inlet and outlet manifolds to the tubes. The brazing step forms a series of brazed joints between the tubes and the manifolds. The method further includes the steps of forming a layer of coating material on exposed portions of the tubes and forming a single piece jacket construction around the tubes.

Abstract: A gas generator or rocket engine includes: a first storage tank configured to contain a high-pressure liquid propellant; a nozzle connected to the first storage tank; a heat source connected between the first storage tank and the nozzle and configured to add heat to the high-pressure liquid propellant at a heat transfer rate to substantially gasify the high-pressure liquid propellant, where the nozzle is configured to expel and expand the substantially gasified high-pressure liquid propellant, and where the gas generator is configured so that an expanded temperature of the substantially gasified high-pressure liquid propellant after being expanded by the nozzle is in the range ?50° C. to 100° C., preferably 0° C. to 50° C.; and a controller connected to the heat source and configured to adjust the heat transfer rate.

Abstract: A wall structure capable of withstanding high thermal load, including a wall structure with an inner wall, an outer wall positioned about the inner wall, and distancing elements located between the inner and outer walls. The distancing elements are joined to at least one of the inner and outer walls by means of laser welding and thereby establishing a plurality of cooling ducts between the inner and outer walls and a weld joint formed by the laser welding and having a cross-section taken through the wall structure that is substantially T-shaped and which displays a rounded shape towards an inside one of the plurality of cooling ducts.

Abstract: An apparatus and method are provided for decomposition of a propellant. The propellant includes an ionic salt and an additional fuel. Means are provided for decomposing a major portion of the ionic salt. Means are provided for combusting the additional fuel and decomposition products of the ionic salt.

Abstract: An actively-cooled, fiber-reinforced ceramic matrix composite thrust chamber for liquid rocket propulsion systems is designed and produced with internal cooling channels. The monocoque tubular structure consists of an inner wall, which is fully integrated to an outer wall via radial coupling webs. Segmented annular void spaces between the inner wall, outer wall and adjoining radial webs form the internal trapezoidal-shaped cooling channel passages of the tubular heat exchanger. The manufacturing method enables producing any general tubular shell geometry ranging from simple cylindrical heat exchanger tubes to complex converging-diverging, Delaval-type nozzle structures with an annular array of internal cooling channels. The manufacturing method allows for transitioning the tubular shell structure from a two-dimensional circular geometry to a three-dimensional rectangular geometry.

Abstract: Device for supplying fuel for a rocket propulsion unit has a first and at least a second fuel circuit for respective different fuels. Each fuel is brought to an increased energy level by a pump and is supplied for combustion by way of injection elements. The first fuel is heated in cooling channels extending in a propulsion chamber wall before the fuel is supplied for combustion, and the first fuel is subsequently fed to at least the turbines assigned to the pumps. A heat exchanger is provided in which the fuel coming from the turbines is in a heat exchange with a fuel coming from a pump. A heat exchanger especially usable in the device for supplying fuel is provided.

Abstract: A method and apparatus are provided for rejecting waste heat (12) from a system (10) including a combustion engine (14), an oxidizer supply (16), and a fuel supply (18). The method includes the steps of providing a fuel flow (32) from the fuel supply (18), providing an oxidizer flow (30) from the oxidizer supply (16), splitting one of the fuel flow and the oxidizer flow into cooling flow (34) and a combustion flow (36), rejecting system waste heat (12) to the cooling flow (34), combusting the combustion flow (36) with the other of the fuel flow (32) and the oxidizer flow (30) to supply combustion gas for the combustion engine (14), and directing the cooling flow (34) into the combustion gas to reduce the temperature of the combustion gas.

Abstract: A space craft's rocket engines are cooled by a recirculating cooling system containing a non-propellant coolant fluid, such as water and/or ethylene glycol. With that recirculating cooling system to maintain the rocket engine combustion chamber at a lower temperature, spacecraft rocket engines may be constructed less expensively and can operate with greater safety by employing the more common metals in their construction. The cooling system also provides an easy means to warm and/or vaporize a propellant.

Abstract: For a liquid rocket engine including the SSME the gaseous heat exchanger that is in the liquid oxygen turbopump exit flow path that serves to gasify the liquid oxygen for pressurizing the liquid oxygen tank and the POGO is modified by changing the flow circuit to replace the liquid oxygen with high pressure hydrogen tapped off of the low pressure fuel turbopump so that hydrogen/hydrogen at a favorable pressure differential is in indirect heat relation for providing a safe environment and avoiding what may result in a catastrophic failure in the event the heat exchanger fails. An external heat exchanger in communication with the heat exchanger in the liquid oxygen turbopump is provided to gasify the liquid oxygen for the pressurizing purposes.

Abstract: A hybrid heater vaporization system (280, 380) including hybrid heaters (250, 350) can be used to vaporize liquid oxygen as it enters a motor (200, 300) to improve the combustion characteristics of a hybrid rocket (11). The motor (200, 300) preferably includes hybrid fuel both in a substantially cylindrical portion (216, 316) in the substantially cylindrical portion of the motor and a substantially multi-toroidal shaped portion (217, 316) in the forward end of the motor (200, 300). The vaporization system (280, 380) of the present invention finally makes hybrid rockets (11) practical for aerospace applications.

Abstract: Cooling of engine walls with fuel, wherein the walls have a structure which has an inner wall, to which hot gas is admitted during the operation, a colder outer wall, as well as a plurality of webs which connect the walls and divide the hollow space present there into a plurality of cooling ducts.The fuel is introduced in the cold state into the wall structure, is delivered through the cooling ducts while absorbing heat via the inner wall, and is subsequently used to generate thrust.A hot fluid flow taken from the engine is admitted during the operation from the outside to the "cold" outer wall of the wall structure, through which fuel flows, and this "cold" outer wall thermally expands as a result, or its thermal contraction is reduced.

Abstract: A fuel member comprises a continuous, elongated, flexible and combustible structural component that has a length to width of ratio greater than 100:1, and contains within the structural component, either as a part thereof or as a separate component, an oxidizer in an amount sufficient to support combustion of the fuel member. The fuel member is essentially self-supporting in that it does not require containment tanks or rigid encasement, and therefore is particularly suitable for use in single stage, reusable rocket propulsion systems.

Abstract: In the preferred embodiment of the invention, a cooled bipropellant thruster 70 for controlling the on-orbit position and orientation of a spacecraft is provided. The cooled bipropellant thruster 70 uses a liquid fuel and liquid oxidizer. The liquid fuel is decomposed in a first chamber 72 with a catalytic bed of decomposition material 74 and produces at least one reaction gas, which flows to the second reaction chamber 90. The second reaction chamber 90 is heated by the reaction gas, but is cooled by liquid oxidizer flowing through cooling passages 92, 94, 98, and 102, which brings the oxidizer into a heat exchange relationship with the second reaction chamber 90. During the heat exchange relationship, heat is transferred from the second reaction chamber 90 to the oxidizer and the oxidizer transforms into a gas. The gaseous oxidizer is fed into a second reaction chamber 90 to secondarily react with the reaction gas.

Abstract: Each of the tubes of the tube thrust chamber of a rocket engine is hooked at either end with the end at the outlet oriented horizontally relative to the rocket engine's center line and the end at the inlet oriented vertically and bent outwardly away from the rocket thrust chamber and facing the outlet for fitting into apertures formed in the outlet and inlet manifolds. The method of construction and assembly is manifested by assembling the sub-assembly of the wedge ring/manifold above the open ends of the vertical portion of the tubes and lowering the assembled wedge ring/manifold so that each tube fits into complementary apertures formed in the inlet manifold and the wedge ring tightly fits onto the outer periphery of the tubes and brazing this assembly in situ. A skirt extends below the thrust chamber and includes a planar surface for resting on ground to support the thrust chamber in an upright position.

Abstract: A propulsion system for use with a missile or like aerial projectile is disclosed which is suitable for use to operate the divert thrusters and the attitude control thrusters of such a missile while using non-toxic propellants which are entirely non-reactive during storage, transportation, and handling. In the preferred embodiment of the present invention, highly refined liquid hydrocarbon fuel and oxygen gas are used as the propellants, with a relatively small amount of the liquid hydrocarbon fuel and a relatively large amount of the oxygen gas being combusted in an oxygen heater to produce a hot oxygen gas containing only small amounts of the products of combustion. The liquid hydrocarbon fuel and the hot oxygen gas are burned in divert thrusters, and, optionally, in attitude control thrusters to produce thrust.

Abstract: In the preferred embodiment of the invention, a cooled bipropellant thruster 70 for controlling the on-orbit position and orientation of a spacecraft is provided. The cooled bipropellant thruster 70 uses a liquid fuel and liquid oxidizer. The liquid fuel is decomposed in a first chamber 72 with a catalytic bed of decomposition material 74 and produces at least one reaction gas, which flows to the second reaction chamber 90. The second reaction chamber 90 is heated by the reaction gas, but is cooled by liquid oxidizer flowing through cooling passages 92, 94, 98, and 102, which brings the oxidizer into a heat exchange relationship with the second reaction chamber 90. During the heat exchange relationship, heat is transferred from the second reaction chamber 90 to the oxidizer and the oxidizer transforms into a gas. The gaseous oxidizer is fed into a second reaction chamber 90 to secondarily react with the reaction gas.

Abstract: An augmentor screech suppressor for a gas turbine engine includes a screech suppressor liner mounted on a screech housing and an actuator for translating the screech suppressor liner axially forward and aft, with the actuator located outside the outer duct of the engine. In a preferred embodiment in a variable cycle gas turbine engine, a variable area bypass injector (VABI) valve with a circumferentially conical surface slopes into the screech suppressor liner surface and inner and outer crank arms transfer linear motion from the actuator through the outer duct wall, cooperating with idler arms to translate the screech suppressor housing, and thus the VABI valve and screech suppressor linear, axially between open and closed valve position.

Abstract: A regulator valve assembly adapted to be used in conjunction with a torque converter in a vehicular transmission having a pressurized hydraulic distribution system, a cooling system and a torque converter. The regulator valve assembly has a stepped valve chamber to receive a spool valve member for axial translation. The spool valve member defines first, second, third and fourth subchambers within the stepped valve chamber. Axial translation of the valve member effects selective communication between the subchambers. A first land on the spool valve member is interposed between the first and second subchambers. A second land is interposed between the second and third subchambers. A third land is interposed between the third and fourth subchambers, and a fourth land is exposed to the fourth subchamber. A first outlet communicates with the first subchamber. A first inlet and a second outlet communicate with the second subchamber. A second inlet and third outlet communicate with the third subchamber.

Abstract: Difficulties in controlling the flow of oxidant to a combustor (10) in a stored energy system for driving a gas turbine (16) resulting from Joule-Thompson cooling of the oxidant are avoided in a system wherein a hot gas distributor (50, 50', 60) is disposed within an oxidant storage vessel (40, 60) to add heat to gases near the outlet (42) of the bottle (40, 60) as a result of the firing of a squib (54) in fluid communication therewith. The heating of the gases at the outlet (42) raises the temperature of the oxidant sufficiently to offset Joule-Thompson cooling.

Abstract: A liquid-solid propulsion system having a tank of liquid oxygen and a high pressure chamber loaded with solid grain fuel with a portion of the liquid oxygen being passed through a heat exchanger to convert the liquid oxygen to gaseous oxygen. The gaseous oxygen is directed to the chamber or solid grain fuel to induce a fuel rich gas burn that is directed to a thrust chamber which also receives liquid oxygen to increase the characteristic velocity of the exhaust and thereby provide the specific impulse of the propulsion system. The gaseous oxygen is also directed to the liquid oxygen tank to pressurize the flow of liquid oxygen from the tank. Valves are interposed to control the flow of liquid and gaseous oxygen to provide the required mixture ratio in the thrust chamber for optimum specific impulse or to terminate the thrust of the propulsion system. A method for providing the liquid-solid propulsion system is also disclosed.

Abstract: In a heat exchanger system wherein a primary conditioning fluid is reactive with a fluid to be conditioned, a leak safe arrangement of concentric tubes disposed across the conditioned fluid flow path having the reactive primary fluid in a central duct surrounded by inert fluid in an outer duct so that single failure leak of primary conditioning fluid or inert fluid presents no risk of harmful reaction with conditioned fluid. The heat exchanger system is used in an air collection and enrichment system in a hypersmic vehicle propulsion system.

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 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 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: An inner tube tube an outer surface of which is machined to define regerative coolant grooves for flowing a coolant to cool a combustor is split at a throat into upper and lower portions which in turn are fitted into an outer tube fabricated as a unitary structure. The upper and lower inner tube portions and the inner and outer tubes are integrally joined together.

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: An inlet air demoisturizing system utilizes a plurality of screens between the inlet of a cryogenic engine and a heat exchanger therefor. In a first embodiment, a curtain of liquid air flows on the forward side of the screen to pre-cool the gaseous intake air below the freezing point of water. In a second embodiment, a mist of liquid air is used to pre-cool the air prior to arrival at the screen. The screen traps ice crystals precipitated from the air. In each embodiment, a second sprayer washes the screen with a curtain of liquid water to remove the ice crystals therefrom. Collector troughs collect the liquid air and suspended ice crystals for removal from the engine inlet.

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 propulsion system and method having means for cooling the combustor liner and throat liner of a rocket casing by endothermic pyrolysis of a hydrocarbon fuel in the fuel passageway which is adjacent to and surrounding the combustor liner and the throat liner. Means is provided for high heat flux to the combustor liner and throat liner from combustion within the rocket casing whereby the temperature of the liners exceeds their thermal limits. Catalyst means are utilized for the endothermic pyrolysis. Hydrocarbon fuel is passed through the fuel passageway; heat is provided from the combustion of fuel in the combustion chamber to the fuel passageway by radiation through the combustor liner and the throat liner; and the hydrocarbon fuel is heated at a temperature sufficient to cause the endothermic pyrolysis of the hydrocarbon in the fuel passageway.

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: A propulsion system which utilizes hydrocarbon and hydrogen fuel, having means for cooling of the combustor liner and throat liner of a rocket casing by endothermic pyrolysis of the hydrocarbon in the presence of hydrogen in the fuel passageway. The hydrogen in the fuel accelerates the rate of endothermic pyrolysis in the fuel passageway which is adjacent to and surrounds the combustor liner and the throat liner. Means is also provided for high heat flux to the combustor liner and throat liner from combustion within the rocket casing so that the temperature of the liners exceeds their thermal limits. By the propulsion system and method, hydrocarbon fuel and hydrogen are passed through the fuel passageway which is adjacent to and surrounds the combustor liner and the throat liner in a rocket casing, and heat from the combustion of fuel and hydrogen in the combustion chamber is provided to the fuel passageway by radiation through the combustor liner and the throat liner.

Abstract: A thruster for use in a spacecraft uses vaporized propellant in its combustion chamber (110). Normally two propellants are used of the kind used for main thrusters so that two vaporization chambers are needed. Pressure valves (108, 109) are controlled by electronic means (220) to maintain current pressure.

Abstract: A thruster system has a main bipropellant propulsion unit including a propulsion engine (50) and bipropellant supply having a tank (3) for an oxidizing agent and a tank (4) for a fuel. An auxiliary cold propellant thruster (21) uses propellant supplied by bleeding off vapor from the contents of the oxidizing agent tank (3).

Abstract: An air liquification system for a combustor or the like, including a condenser having an air intake and a heat exchanger for cooling the air to substantially a liquid state. A first conduit conducts the liquid air to the combustor. A source of hydrogen in a liquid state is provided, and a second conduit conducts the liquid hydrogen to the heat exchanger whereby the hydrogen acts as the cooling medium therein. A third conduit conducts the hydrogen from the heat exchanger to the combustor and includes a compressor for compressing the hydrogen up to a given combustor pressure. A second combustor may be provided for operating at ambient air pressure, with a portion of the hydrogen from the condenser being conducted to the second combustor upstream of the compressor. An air precooler may be provided upstream of the heat exchanger for passing the liquid air therethrough to act as a regenerative cooling medium therein prior to conducting the air to the first combustor.

Abstract: The present invention provides a regeneratively cooled rocket combustor comprising a combustor wall, a plurality of coolant channels extending through the combustor wall and at least one keel-rib extending into each coolant channel from the channel roof and which is bounded on both of its sides by arcuate flutes for providing smooth and continuous surface transition between the keel-rib and adjoining interior surfaces of the coolant channel.

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

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: In the cooling wall of a regenerative cooling rocket combustion chamber, the throat portion thereof in which the heat load is highest is formed of a porous material having a perspiring cooling action. The other portion in which the heat load is relatively low is designed to be cooled by conventional forced convection. By such a composite cooling system, a high cooling effect is provided and yet the pressure loss in the cooling passages is small and light weight and long life of the engine are attained concurrently.

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

Abstract: In a heat exchanger wall construction, such as used in regeneratively cooled combustion chambers for liquid fueled rocket engines, the wall construction consists of an inner wall with cooling channels spaced apart by webs. The cooling channels are open on one surface of the inner wall. An outer wall contacts the webs of the inner wall and forms a closure over the openings from the cooling channels. The outer wall consists of an intermediate layer electro-deposited on the surface of the inner wall in which the cooling channels are formed so that it provides a closure surface for the channels. The surfaces of the channels within the inner wall are coated by electro-depositing a layer of gold on them. After forming the gold layer on the surfaces of the cooling channels, the outer surfaces of the webs are etched so that the gold layer extends outwardly from the outer surface of the webs. Next the intermediate layer is deposited and it can be either a single layer or multi-layered.