Abstract: An acoustic resonance igniter uses high-pressure helium to heat a resonance cavity so a hot surface of the resonance cavity forms a source of ignition to a combustion chamber. The resonance cavity may be round or may extend linearly to increase the size of the hot surface. The combustion chamber is cooled by arranging a feed of hydrogen and oxygen which is oxygen rich and which becomes more so when ignition occurs. A second combustion chamber receives the combustion chamber output and adds additional hydrogen through ports tangential to the wall of the second combustion chamber to enrich the fuel ratio and cool the second combustion chamber. The acoustic resonance igniter is used to ignite a large rocket engine or to form a rocket thruster.

Abstract: An ignition system and method of igniting the ignition system includes a main catalyst section in a staged relationship with a pilot-catalyst section to stage a decomposition though the pilot-catalyst section which preheats the main catalyst section.

Abstract: One embodiment is directed to a method for testing an aircraft prior to flight. The method includes receiving a user signal from a pre-flight test input source, the user signal indicating that a pilot of the aircraft has directed engine control circuitry, which is arranged to electronically control operation of a set of piston engines of the aircraft during flight, to begin testing the aircraft in an automated manner. The method includes, in response to the user signal, conducting a pre-flight test of the aircraft from the engine control circuitry. The method includes, upon completion of the pre-flight test, outputting a result of the pre-flight test from the engine control circuitry.

Abstract: A propulsion system for creating a propulsive force has a combustion chamber, a pair of electrodes within the combustion chamber and a power supply attached to the electrodes to create a high voltage field within an initiation zone of the combustion chamber, and an injector for introducing a propellant, preferably in atomized form, into the high voltage field for creating the propulsive force. In a preferred embodiment of the present invention, the propellant is hydrogen peroxide. In another embodiment of the present invention, a second propellant is introduced into the combustion chamber for increasing the propulsive force.

Abstract: A bi-propellant injector (66) includes a first injector element (68) and a second injector element (70) injecting a first propellant (69) and a second propellant (71), respectively, into a combustion chamber (53). A flame-holding zone igniter (74) is adjacent to and ignites recirculation of at least a portion of the first propellant (69) and at least a portion of the second propellant (71) within a flame-holding zone (76).

Abstract: The reduced toxicity hypergolic bipropellant fuels of the present invention contain a hypergolic fuel and rocket grade hydrogen peroxide oxidizer, preferably HTP. The hypergolic fuel may be a reactive fuel or a catalytic fuel. The rocket grade hydrogen peroxide oxidizer consists of about 90 weight % to about 99 weight % H2O2, more preferably about 94 weight % to 99 weight % H2O2. However, hypergolic ignition may be attained with a H2O2 content as low as 70 weight % for some reactive fuels. The reactive fuel comprises about 6 weight % to 10 weight % reducing agent in a solvent. The catalytic fuel comprises about 6 weight % to 10 weight % catalytic agent in a solvent. The reactive fuels contain ingredients that are intrinsically reactive with rocket grade hydrogen peroxide. Upon contact with rocket grade hydrogen peroxide oxidizer, reactive fuels react vigorously with the hydrogen peroxide resulting in ignition. The catalytic fuels are produced by dissolving a catalytic agent in a solvent.

Abstract: The non toxic bipropellent of the present invention contains a non-toxic hypergolic miscible fuel (hereinafter referred to as “NHMF”) and a rocket grade hydrogen peroxide. This NHMF has rapid ignition capability and minimizes the formation of precipitate, even when exposed to extreme heat or water. The NHMF of this invention contains 5 species. Namely, a polar organic species miscible with hydrogen peroxide; a propagator, which may be substituted or unsubstituted amines, amides or diamines; an inorganic metal salt, which reacts to form a catalyst in solution or as a colloid; acetic acid; and alkali acetate. The inorganic metal salt is miscible with the polar organic species and the propagator in solution.

Abstract: The invention concerns an acoustic igniter for igniting a mixture of rocket fuels in a liquid propellant rocket engine combustion chamber comprising a cylindrical precombustion chamber (101) including a cylindrical wall (111) and first and second end walls (112, 113), a rocket fuel injection nozzle (103) emerging into the precombustion chamber (101) through the first end wall (112) via an orifice of diameter dn, a rocket fuel injector (104) arranged inside said nozzle (103) along the axis thereof, at least an outlet orifice (102) of minimum diameter df provided in the cylindrical wall (111), an acoustic resonator (105) defining a cavity opening into the precombustion chamber (101) opposite the nozzle (103), through the second end wall (113), via an orifice (151) of diameter dr. The acoustic resonator (105) is enclosed by a housing (106) which defines around the acoustic resonator (105) a closed auxiliary chamber (160) which communicates only with the precombustion chamber (101) by at least one conduit (107).

Abstract: The invention concerns an acoustic igniter for igniting a mixture of rocket fuels in a liquid propellant rocket engine combustion chamber comprising a cylindrical precombustion chamber (101) including a cylindrical wall (111) and first and second end walls (112, 113), a rocket fuel injection nozzle (103) emerging into the precombustion chamber (101) through the first end wall (112) via an orifice of diameter dn, a rocket fuel injector (104) arranged inside said nozzle (103) along the axis thereof, at least an outlet orifice (102) of minimum diameter df provided in the cylindrical wall (111), an acoustic resonator (105) defining a cavity opening into the precombustion chamber (101) opposite the nozzle (103), through the second end wall (113), via an orifice (151) of diameter dr. The acoustic resonator (105) is enclosed by a housing (106) which defines around the acoustic resonator (105) a closed auxiliary chamber (160) which communicates only with the precombustion chamber (101) by at least one conduit (107).

Abstract: The non-toxic bipropellent of the present invention contains a non-toxic ergolic miscible fuel (NHMF) and a rocket grade hydrogen peroxide. This non-toxic hypergolic miscible fuel (NHMF) has rapid ignition capability. The non-toxic hypergolic miscible fuel (NHMF) contains 3 species. Namely, a polar organic species miscible with hydrogen peroxide, a propagator, which may be substituted or unsubstituted amines, amides or diamines, and an inorganic metal salt, which reacts to form a catalyst in solution or as a colloid. The inorganic metal salt is miscible with the polar organic species and the propagator in solution. The catalyst has a faster rate of reaction with said rocket grade hydrogen peroxide than the propagator, the propagator has a faster rate of reaction with the rocket grade hydrogen peroxide than the polar organic species, and the polar organic species, propagator and catalyst are mutually soluble.

Abstract: The multifuel combustion engine is a Lenoir cycle (constant volume), pulst combustion engine capable of operating on gasoline, diesel or kerosene based fuels. Although the preferred embodiment is described in terms of Lenoir cycle pulsejet engines, the present invention has application to combustion engines in general. The conventional Lenoir cycle engine has been modified to provide a direct, premixed fuel-air spray to the combustion chamber and means for igniting the fuel-air spray. Said fuel-air spray is separate and distinct from the fuel-air charge which is fed to the combustion chamber from the engine head. Additionally, means are provided for preheating the combustion chamber so that the same fuel-air ratio mixes can be fed to the combustion chamber for cold start or hot restart of the engine. A method of burning different fuels in combustion engines is also claimed. The present invention includes the application of the modified Lenoir cycle engines to smoke generator equipment.

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: In a propulsion system which includes a rocket casing having a combustor chamber with a combustor liner and a nozzle throat chamber with a nozzle throat liner; a fuel passageway adjacent to and surrounding the combustor liner and nozzle throat liner which form the inner wall of the fuel passageway, the outerwall of the fuel passageway being spaced from the innerwall to form the passageway, a plurality of flow directing vances disposed in the fuel passageway to direct the fuel in the fuel passageway circumferentially. In the propulsion system, the fuel travels a longer path through the passageway, and the residence time of fuel in the passageway is increased to promote heat transfer to the fuel in the passageway. A plurality of fuel injection holes in the combustor liner are also arranged so that fuel is injected into the combustion chamber in a direction which promotes circumferential motion of the fuel.

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: Disclosed is a method of converting non-hypergolic, liquid, rocket propellants into hypergolic propellants. The method to accomplish the conversion relates to the use of ammonium metavanadate as an additive to the liquid oxidant, such as, red fuming nitric acid (RFNA). The RFNA with additive is hypergolic with the usual fuels with which RFNA has been employed where a separate ignition system is normally required. RFNA with additive is also hypergolic with fuels which have been non-hypergolic and which have not been so used in the past with RFNA. These additional fuel blends include a fuel material selected from turpentine, aniline, triethylamine, furfuryl alcohol or blends of these fuel materials.