SE464719B - Propulsion system - Google Patents

Description

464 719 2 to eliminate the requirement for vacuum-type storage tanks and other bulky systems for storing liquid fuels in aircraft.

Mixtures of hydrocarbons and liquid oxygen have been widely used as propellants in aircraft propulsion systems. It is known that such hydrocarbon fuels burn in combustion chambers at temperatures exceeding 27,5006 (5,000 ° F) and generate a large thrust. However, propulsion systems for newer aircraft designs require greater draft power than is usually achieved by the hydrocarbon-liquid nitrogen mixtures, and more complex fuels in more complex propulsion systems have been developed to provide higher traction. Most of these complex fuels cannot be used in conventional propulsion systems for various reasons, such as extremely high combustion temperatures, and complex propulsion systems and storage systems must be developed to handle the combustion of such fuels. The complex propulsion systems that overcome these difficulties contribute significantly to the weight of the aircraft. From the above it appears that it would be advantageous to be able to use less complex propulsion systems based on hydrocarbon-liquid oxygen mixtures in order to overcome the above-mentioned disadvantages.

It is well known that certain chemical substances such as high molecular weight hydrocarbons react endothermically to produce reaction products which can be used as fuel. Previous systems have been developed which convert certain chemical substances by endothermic reactors into fuels which can be combusted in a burner. However, many of these prior systems are disadvantageous because the endothermic reaction process generates certain reaction products, which can build up an undesirable coating in the burner. It is therefore desirable to provide a propulsion system which utilizes the simple hydrocarbons, such as ethylene, ethane, propylene, and propane, as the chemicals which react endothermically to produce preferably only reaction products useful as fuels in propulsion systems for aircraft.

In previous systems which use fuels which generate high temperatures in the propulsion system, such as in a combustion chamber, it has been difficult to find materials which can withstand the high temperatures, ie. 2750 ° C (5000 ° F) or more. Many types of burner liners and neck liners of the rocket casing have been proposed, which, however, are complex and expensive or have a limited life or require circulation thereby of large amounts of coolant to prevent weakening and / or melting of the liner material. Accordingly, it is desirable to provide improved burner and neck constructions and materials in propulsion systems which overcome the above-mentioned disadvantages. 464 719 Summary of the Invention.

Accordingly, a principal object of the present invention is to provide an aircraft and propulsion system of the type which utilizes multiple propellants.

Another object of the invention is to provide a propulsion system and method which utilizes a selection of propellants in a systematic manner to achieve flight to orbit in one step taking into account the requirements of both the aircraft and propulsion system design.

Another object of the present invention is to provide a fuel storage system and method for storing several fuels in a light and simple vessel construction, and providing a maximum of fuel reserve in an aircraft of minimal size and weight.

Another object of the invention is to provide a fuel storage device and method which provides maximum fuel reserves while providing optimum aircraft shape with respect to lift capacity and air friction.

A further object of the present invention is to provide a propulsion system and method which utilizes burner liners and nozzle neck liners to maintain high combustion pressures and which can withstand high combustion temperatures.

A further object of the present invention is to provide a propulsion system and method which utilizes a mixture of hydrocarbon and hydrogen gas fuel with a liquid oxygen oxidizer to achieve high combustion temperatures from low molecular weight fuels. Yet another object of the invention is to provide a propulsion system and method for cooling the liners of the torch and nozzle neck of a rocket engine.

Another object of the present invention is to provide an improved propulsion system and method in which the burner liner and nozzle neck liner of a rocket engine are cooled by the endothermic decomposition of the fuel.

A further object of the invention is to provide an improved rocket burner and nozzle neck construction which increases heat transfer and residence time for endothermic cooling of the burner liner and nozzle neck liner.

A further object of the present invention is to provide a fuel injection system and liquid oxygen injection system for injecting fuel and liquid oxygen into the burner chamber of a propulsion system. These and other objects are achieved in that the propulsion system according to the present invention has obtained the features stated in claim 1.

The propulsion system of the present invention utilizes hydrocarbon fuel and includes a rocket housing having a burner liner and a neck liner, a hydrocarbon fuel fuel passage located adjacent to and surrounding the burner liner and liner, means for providing fuel flow in means of providing a fuel passage; large heat flow to the burner liner and neck liner from combustion within the rocket casing, with the liner temperature exceeding their thermal limitations, as well as means for cooling the burner liner and neck liner by endothermic pyrolysis of the hydrocarbon in the fuel passage. In one embodiment, the endothermic pyrolysis of hydrocarbon is performed in the fuel passage in the presence of hydrogen, and the hydrogen in the fuel accelerates the rate of the endothermic pyrolysis. According to another embodiment, the endothermic pyrolysis of hydrocarbon is carried out in the fuel passage in the presence of catalysts which accelerate the rate of the endothermic pyrolysis. The endothermic pyrolysis of the hydrocarbon fuels used in the present invention, in the presence of hydrogen and / or catalysts, yields a fuel product with higher combustion rates, a fuel product with higher combustion temperatures and / or a fuel product with a lower molecular weight.

The aircraft according to the present invention has a propulsion system which utilizes a dual fuel system, where at least one of the fuels is a cryogenic fuel. An inner tank without vacuum contains the cryogenic fuel and an outer tank without vacuum, which surrounds the inner tank, contains a second fuel, which is a fuel with a low freezing point and a high boiling point, which acts as an insulator for the cryogenic fuel in the inner tank. The other fuel in the outer tank can also be a cryogenic fuel. The fuel tank or vessel system comprises a pressurized cylindrical inner vessel or tank for containing or storing the liquid cryogenic fuel, e.g. hydrogen, and has a rigid insulated wall. The inner tank for containing and storing liquid cryogenic fuel is surrounded by an outer tank for containing the second fuel, which is generally a liquid and / or gaseous hydrocarbon. By surrounding the liquid hydrogen containing, for example, the inner tank with pre-cooled liquid and / or gaseous hydrocarbon, excessive boiling of the liquid hydrogen at high altitudes is prevented.

The liquid and / or gaseous hydrocarbon layer insulates the liquid hydrogen, thus avoiding the need for conventional cryogenic vacuum vessels as insulation.

The wall of the outer hydrocarbon tank forms the shell of the aircraft. The outer wall of the outer tank is thus adapted to the aerodynamic shape of the aircraft. The 5,464,719 pre-cooled hydrocarbon is substantially uncompressed and therefore fills the entire space between the liquid hydrogen tank and the shell of the aircraft. This design and method for storing fuels in the aircraft allows the design of the aircraft for optimal lifting capacity and air resistance without reducing fuel reserves.

Liquid oxidizer storage means are provided, and conventional means are used to keep the liquid oxidant in the subcooled state to reduce boiling. The liquid oxidant storage means may be suitably arranged to fill a non-pressurized vessel or several vessels suitable for the shape of the aircraft and the improved fuel storage system of the present invention.

According to the present invention there is also provided a propulsion system having a rocket housing containing a burner, fuel injectors, nozzle neck and nozzle, a fuel passage in which fuel is pyrolyzed, and which is adjacent to and surrounds the burner and nozzle neck, and means for providing fuel in fuel. wherein the improvement consists of an inner wall woven of silicon carbide fibers, which forms a burner liner and a nozzle neck lining, and an outer wall woven of silicon carbide fibers, which is separated from the inner wall for forming the fuel passage. The silicon carbide fibers are woven, preferably continuously, and conduct heat from the burner and nozzle neck to the fuel passage, thereby providing heat for the endothermic pyrolysis of fuel. In preferred embodiments, the inner wall of woven silicon carbide fibers has a porosity for selectively controlling diffusion of hydrogen from the passage through the wall into the burner and nozzle neck. This improvement of the propulsion system, i.e. in rocket engines, where the combustion chamber is made of continuously woven silicon carbide fibers, allows the operation of rocket engines at very high temperatures by means of low molecular weight fuels, ie. for example fuels such as low molecular weight hydrocarbons and hydrogen in the vicinity of the oxidizer, oxygen. The combustion products from these fuels have a relatively low molecular weight. The low molecular weight hydrocarbons used herein are unsaturated or saturated hydrocarbons having less than 4 carbon atoms.

Hydrogen available from the liquid hydrogen in the hydrocarbon mixture in addition to the hydrogen obtained from the endothermic decomposition of hydrocarbon in the fuel passage promotes the formation of a large fraction of water vapor with a relatively small molecular weight in the combustion products. High combustion temperature in the burner, and the high temperature in the nozzle neck, are allowed by the combined use of continuously wound silicon carbide fibers and by film cooling from the hydrogen in the fuel passage, where the hydrogen diffuses from the fuel passage through the burner liner into the burner liner and neck liner. nozzle neck chambers.

The net result is a ratio between combustion temperature and molecular weight, which gives a high speed of sound in the nozzle neck of the rocket engine and after supersonic expansion, a high blowout speed of the rocket engine.

According to the present invention, the woven silicon carbide fiber structure of the burner liner and neck liner provides high ring strength, and the fibrous layers provide sweat cooling through the controlled porosity of the wound silicon carbide fibers, which allows migration or diffusion of hydrogen, but not hydrocarbon, into pressurized hydrogen. and the nozzle neck chamber. Film cooling of the hot side of the wall is thus achieved while at the same time the supply of hydrogen reduces the molecular weight of the exhaust gases, which gives a higher neck velocity. In addition, the hydrogen film on the burner and nozzle neck side of the burner liner and the nozzle neck liner reduces fluid wall friction in the nozzle neck.

According to another aspect of the present invention, the propulsion system includes a rocket housing having a burner chamber with a burner liner and a nozzle neck chamber with a nozzle neck liner, a fuel passage adjacent and surrounding the burner liner and the nozzle neck liner forming the inner wall of the inner wall passageway. the passage, means for providing fuel flow in the fuel passage, and a plurality of flow directing guide rails arranged in the fuel passage for directing the fuel in the fuel passage peripherally. The fuel thus moves a longer path through the passage, which increases the residence time of the fuel in the passage to promote heat transfer to the fluid in the fuel passage. Due to the flow-directing guide rails, there is a longer fluid residence time in the fuel passage, and this results in higher heat transfer and extended time for the endothermic pyrolysis or decomposition of the hydrocarbon fuel.

In a further embodiment, a plurality of fuel injection holes in the burner liner inject fuel into the combustion chamber in a direction which promotes peripheral movement of the fuel in the combustion chamber. An oxidizer is also inserted through oxidizer injection holes to the center of the combustion chamber to inject oxidizer into the combustion chamber in a direction that promotes peripheral movement of the oxidizer. In this way, the fuel and oxidizer in the chamber are mixed before and during combustion.

The present invention uses hydrocarbon fuel to generate fuel products which have at least one of the following properties: (1) higher combustion rates, (2) higher combustion temperatures and (3) lower molecular weight prior to combustion in the combustion chamber. Acetylene is, for example, a preferred fuel product which has a higher combustion rate, a higher combustion temperature and a lower molecular weight than the hydrocarbon from which it is generated.

In one embodiment of the invention, ethylene is thermally converted to acetylene and hydrogen using hydrogen in the fuel passage as a catalyst and / or using an additional catalyst, while excess hydrogen from the fuel mixture also acts to buffer and prevent carbon polymerization, resulting in coke or in some cases, acetylenetonation.

By means of the present invention it is possible to emphasize the consumption of liquid oxygen, liquid and / or gaseous hydrocarbon as well as liquid hydrogen in this order. This is significant due to the fact that at all rocket blowout speeds traction is achieved through fuel mass supply for impulse increase. Liquid oxygen, followed by liquid hydrocarbon, has a high liquid density, and therefore a high proportion of fuel relative to the size and fixed weight of the aircraft.

The propulsion systems of the present invention can be used in any aircraft in which the engine or engines use a rocket casing. For example, the invention may form part of a rocket motor, frame motor, or scram motor.

Brief description of the drawings.

The invention will now be described in more detail below with reference to the accompanying drawings, in which: Fig. 1 is a partially isometric perspective view of a rocket-propelled aircraft showing the propellant vessels of the present invention; Fig. 2 is a sectional view taken along line 2-2 of Fig. 1; the fuel and oxidizer tanks of the aircraft, Fig. 3 is a cross-sectional view of a segment of the fuel tanks showing the construction of the walls of the tanks, Fig. 4 is a cross-sectional view of the upper part of a rocket engine and shows the fuel passage, neck nozzle liner, combustion liner and oxidizer liner. Fig. 5 is a cross-sectional view of the rocket engine taken along lines 5-5 of Fig. 4 and further showing the fuel flow in the fuel passage and flow control guide rails of the present invention; Fig. 6 shows a cross-sectional view of the combustor liner of the present invention.

Detailed description of the invention.

Fig. 1 shows a typical aircraft 10 comprising a cabin and / or payload space 24 in the front part of the aircraft, three engines 8 and a tail section 22 in the rear part of the aircraft, and the improved fuel storage system according to the present invention comprising a tank 4 for liquid hydrogen surrounded by a hydrocarbon tank 20 as well as tanks 6 for liquid oxygen located in the middle to rear sections of the aircraft. All aircraft designs can be used in conjunction with the improved propulsion system and methods of the present invention.

Referring to Figures 1 and 2, liquid and / or semi-solid hydrogen is entrapped and pressurized in the liquid hydrogen tank 4 of the rocket-propelled aircraft or aircraft 10. Semi-solid hydrogen is a combination of liquid and solid hydrogen.

The hydrocarbon tank 20 surrounds the liquid hydrogen tank 4. Here, the liquid hydrogen tank 4 is used in the form of a non-vacuum, inner tank for entraining cryogenic fuel as hydrogen, which normally requires a cryogenic vacuum insulating chamber or vessel, and the hydrocarbon tank 20 is in the form of a non-vacuum, outer tank. which surrounds the inner tank for enclosing a second fuel, such as a fuel having a low freezing point and a high boiling point, which acts as an insulator for the cryogenic fuel in the inner tank 4.

Examples of fuels with a low freezing point are those which have a freezing point in the range from about 150 ° K to about 200 ° K and a high boiling point, such as from about 275 ° K to about 425 ° K. Hydrocarbon fuels, whether they are liquid or gaseous squeegee, and if they are located in the above-mentioned freezing and boiling point areas, provide the necessary thermal barrier to prevent hydrogen decoction. The insulation in the wall separating the liquid hydrogen tank 4 and the hydrocarbon tank 20 is sufficient to prevent the freezing of hydrocarbon by liquid hydrogen for the period of time normally required for flight missions. At a tank pressure of 31 kp / mm 2 (20 psia), liquid hydrogen at a temperature of 36 ° K absorbs 2.8 kcal / kg (5 BTU per pound), and semi-solid hydrogen at a temperature of 25 ° K absorbs 14 kcal / kg (25 BTU per pound) before boiling occurs.

In addition to low freezing point, the hydrocarbon must have a high boiling point (low vapor pressure) so that it does not require compaction at high altitudes. According to the present invention, a compression of less than about 3 kp / mm 2 (2 p.s.i.a.) is acceptable. The low vapor pressure of the hydrocarbon fuel allows the desired design of the hydrocarbon tank 20.

Accordingly, the hydrocarbon tank 20 can be shaped so that the aircraft 20 receives reduced aerodynamic friction and reduced atmospheric frictional heating.

Hydrocarbon fuels which can be used according to the above parameters to provide the necessary insulation for liquid hydrogen include ethylene, ethane, propylene, propane and mixtures thereof. Since these fuels are also cryogenic fuels, the fuel in the outer tank 20 can likewise be a cryogenic fuel. 9 464 71.9 The following table shows examples of typical hydrocarbon fuels that can be used in the hydrocarbon tank 20 of the present invention.

TABLE Hydrocarbon fuels Boiling point Freezing point Liquid density at (° K) (° K) 2400 K kg / m3 Ethylene 305 188 608 CZH4 Ethane 333 182 608 C2H6 Propylene 406 159 640 C3H6 Propane 416 150 624 C3Hs' The fuels shown in the table above can be reduced. desired to reduce their vapor pressure and avoid boiling at high altitudes. The fuels in the table above can, for example, be subcooled to 240 ° K. The liquid density at 240 ° K is shown for each of the hydrocarbon fuels in the table.

The liquid liquid inner tank 4 is substantially cylindrical, as shown in Figures 1 and 2. In preferred embodiments, the tank 4 is an elongate cylindrical tank extending substantially from the cabin and cargo section 24 of the aircraft 10 along the entire length of the aircraft to its rocket section.

For example, the tank 4 is located substantially centrally in the aircraft, as shown in Fig. 2 and extends from approximately line 26 to line 28 in the aircraft 10 as shown in Fig. 1. A liquid hydrogen fuel line 16 is connected to a conventional pump. 18 for liquid hydrogen for supplying liquid hydrogen to suitable lines and a door for distribution to the fuel lines and passages in the propulsion system of the present invention. Although only a fuel line and a liquid hydrogen pump are shown in Fig. 1, those skilled in the art will recognize that a desired number of such fuel lines and pumps may be used, as well as associated auxiliary equipment and control means for the distribution of the liquid hydrogen. 464 719 1 ° The liquid hydrogen tank 4 can be constructed of suitable insulating materials, in which a reasonable thickness prevents liquid hydrogen from freezing the hydrocarbon in the tank 20 for a period of time sufficient for flight missions, As shown in Fig. 3. comprises a typical lightweight wall construction of the tank 4 for liquid hydrogen graphite-epoxy plastic walls 30, which are preferably fiber-reinforced, e.g. with graphite fiber, and separated by an insulating material 40. The insulating material may be, for example, about 25 mm of liquid polyurethane foam injected into and cured between the graphite epoxy plastic walls 30 to form the polyurethane insulation 40. Other typical insulating materials which can withstand the low temperatures include, for example, compacted quartz. The insulating material 40 prevents freezing of the liquid hydrocarbon in the tank 20 by the liquid hydrogen in the tank 4. The inner surface of the liquid hydrogen tank 4 is preferably coated with a material which prevents diffusion of hydrogen through the walls 38. Thus, for example a metal foil suitable for the coating 42 on the inner wall of the tank 4, which comprises an iron-nickel alloy containing about 40 - about 50% nickel and having a low coefficient of thermal expansion. A commercially available foil, which can be used for this purpose, is known in the United States under the brand name INVAR.

According to the present invention, the outer wall of the outer tank, i.e. the outer wall of the outer tank 20 of the shell 30 of the aircraft 10. The outer wall of the outer tank 20 thus adapts to the aerodynamic shape of the aircraft 10. The shell 30 of the aircraft 10, which represents the outer wall of the hydrocarbon tank 20, is preferably about 25 mm thick and is in the form of a so-called "isogrid". Typical shell materials are well known in the art. High temperature materials (about 870 ° C) such as titanium aluminide or some nickel-based heat-resistant alloys, such as Rene 41, can be used as the aircraft shell. The outer shell 30 is bonded to a high temperature type insulating material 36, such as the Min-K commercially available in the United States, which is a thermal insulation reinforced with fibrous material and very fine, heat resistant particulate material and has a microporous structure with very low thermal conductivity and low heat diffusion capacity. Alternatively, for example, compacted or pressed quartz, compacted quartz containing fibrous material and very fine, heat-resistant particulate matter, with a microporous structure with low thermal conductivity and low thermal diffusion capacity, can also be used as high temperature insulation material 36. The inner surface of the tank 20 is the outer wall. also coated with a suitable material to reduce radiation. For example, an iron-nickel alloy containing about 40% to about 50% nickel in the form of a film can be used as the radiation barrier 34 and the coating 34 on the wall of the hydrocarbon tank 20. The foil 34 laminated with the high temperature insulation 36 likewise reduces radiation. The metal foil 34 is also used to reduce heat expansion gradients in the tank.

To avoid freezing of moisture on the outer shell 30 of the aircraft 10, purge gas is used in the spaces 32 of the isogrid structure 31. Hot gaseous nitrogen can, for example, be passed as purge gas through the spaces 32 of the isogrid structure 31 to heat the outer shell 30.

According to the present invention there is provided a method of storing fuels in an aircraft having a propulsion system utilizing a dual fuel system, one of the fuels being a cryogenic fuel which normally requires cryogenic vacuum insulation. The cryogenic fuel, which normally requires a cryogenic vacuum vessel, is placed in an inner, non-vacuum tank and a second low freezing point and high boiling point fuel is placed in the area surrounding the inner tank, the second fuel being an insulator for the cryogenic fuel in the inner tank.

Referring to Fig. 1, the hydrocarbon fuel line 17 is connected to the hydrocarbon fuel pump 15 for supplying hydrocarbon fuel to the correct lines and manifolds according to the invention. The liquid oxygen fuel line 12 supplies liquid oxygen to suitable lines and manifolds via a liquid oxygen fuel pump 14. Conventional fuel supply systems including fuel lines, fuel pumps and fuel manifolds and several sets thereof, as well as control devices and auxiliary equipment (not shown) can be used to supply fuels and oxidizer to the propulsion system of the present invention.

Fig. 4 shows a cross-sectional view of a typical rocket casing 8, which is designed to utilize the fuels proposed according to the present invention.

The rocket housing 8 can form part of a rocket motor, of a frame motor or of a scram motor. A mixture of hydrogen and hydrocarbon, which is supplied from tanks 4 and 4, respectively. 20, cf. Fig. 1, by resp. fuel lines and fuel pumps are led to a manifold 50. The manifold 50 is preferably arranged peripherally around the nozzle of the rocket engine 8 and communicates with the fuel cavity or passage 56 between the fuel wall 54 and the wall 52. The walls 54 and 52 form the nozzle neck 64 and the combustion neck sleeve 62 and The fuel passage 56, which is supplied with fuel from the manifold 50 and in which fuel flows in the direction of the arrows 58, is located adjacent to and surrounds the burner section or chamber shown generally in the area 64 in Fig. 4, and in the area 62 in fig. 4 showed the nozzle neck. In preferred embodiments, the fuel passage 56 surrounds the entire rocket casing so that fuel is supplied to the combustion chamber from the entire periphery of the rocket casing. The fuel passage can have dimensions that are sufficient to allow a sufficient supply of fuel to the combustion chamber and can be easily determined by the person skilled in the art. According to the present invention, the fuel wall or outer wall 54 and the inner wall 52, also called nozzle wall liner 52, nozzle neck liner 52 and burner liner 52, depending on their position in the rocket casing 8, are made of woven silicon carbide fibers. The silicon carbide fibers or threads are woven and conduct heat along the inner wall 52 from the burner 64 and the nozzle neck 62 to the fuel passage 56, thereby providing heat for the endothermic pyrolysis of the fuel in the fuel passage 56. The silicon carbide fibers are well known for high temperatures. substantially in the peripheral direction to provide high pressure containment for the rocket engine. The silicon carbide fibers work up to about 120,000 without cooling.

In preferred embodiments of the present invention, the inner wall 52, i.e. the nozzle neck liner and the burner liner, of a silicon carbide fiber web having such a porosity to achieve selective control of the diffusion of hydrogen from the passage 56 through the wall 52 into the burner chamber 64 and the nozzle neck 62. The outer wall 54 of the fuel passage 54 is made of woven silicon fiber of the high temperature type and is non-porous, so that the hydrogen will not diffuse through the wall 54. The wall 52 is porous so that hydrogen in the fuel passage 56 at a higher pressure. for example about 6300 kp / mm 2 (4000 p.s.i.a.) may pass or diffuse through the wall 52 of the nozzle neck and the burner chamber, which have a smaller pressure than the fuel pressure in the fuel passage 56, e.g. about 4700 kp / mm 2 (3000 p.s.i.a.) This diffusion of hydrogen through the wall 52 into the inner chambers of the rocket casing 3 provides film cooling, as shown by the arrow 90 and the layer 92 in Fig. 6.

As shown in Fig. 6, which shows an enlarged portion of the burner liner 52, hydrogen gas 90 passes or diffuses through the burner liner or nozzle neck liner 52 into the burner chamber or nozzle neck chamber by diffusion through the porous, continuously wound silicon carbide fiber wall 52 to form a film 92. hydrogen on the side of the wall facing the burner chamber and the nozzle neck. The hydrogen diffusion through the wall 52 provides perspiration cooling by migrating hydrogen under pressure through the wall as described above, and is possible due to the porosity of the wall. It is this effect which helps to maintain the temperature of the burner liner and the nozzle neck liner 52 at a temperature which is less than the decomposition temperature of the silicon carbide wires, for example so that the temperature of the wall 52 is kept at less than about 980 ° C to about 1200 ° C.

For the manufacture of the burner liner and nozzle neck liner for a rocket casing comprising a burner, fuel injectors, nozzle neck and some H 464 719, nozzle, a fuel passage in which fuel is endothermally pyrolyzed, and which is adjacent to and surrounds the burner and nozzle to provide fuel flow in the fuel passage, silicon carbide fibers are woven into a plurality of layers in a pattern that promotes high pressure containment of the burner and nozzle neck, and the woven silicon carbide fibers are formed into the shape of a burner and a nozzle neck. The walls 52 and 54 are thus constructed by integrating threads of silicon carbide fibers, preferably in layers, which are circumferentially wound in a continuously woven pattern. The silicon carbide threads or fibers not only promote high pressure containment of the burner and nozzle neck, but are also effective in dissipating heat away from the combustion chamber region 64 and the nozzle neck liner 62 of the rocket casing 8. When winding the silicon carbide threads or fibers around conventional technology, a solid core, easily performed by a person skilled in the art, and the pattern of the fabric can be selected to achieve the porosity required of the wall 52 and the porosity of the wall 54. The wound and shaped silicon carbide fibers can be easily mounted to form the fuel passage 56 with optimal dimensions of within technology well known way.

To achieve the desired porosity to selectively control the diffusion of hydrogen through the woven silicon carbide of the wall 52, various well known methods can be used, e.g. chemical vapor deposition and / or chemical vapor infiltration can be used to deposit or infiltrate the wound or woven silicon carbide fibers with organometallic substances which selectively allow diffusion of hydrogen through the woven silicon carbide.

Referring to Figs. 4 and 5, in the fuel passage 56 a series of flow directing guide rails 80 are located to direct fuel flowing in the fuel passage 56 in the direction of the arrows 58 peripherally so that the fuel moves a longer distance through the passage 56 and thereby increases the fuel flow. residence time in the passage to promote heat transfer to the fluid and / or gas and extend the reaction time for the endothermic pyrolysis of the fuel. Swirl rails 80 bring the gaseous fuel, e.g. hydrogen and hydrocarbons, for rotation for the purpose of promoting heat transfer. The vortex rails 80 may be constructed as an integral part of the outer wall 54, or may be individually mounted on the outer wall 54. The vortex rails may be of a suitable material which can withstand the temperature and pressures of the fuel passage 56, and are preferably made of layers or laminates of the woven, high temperature resistant silicon carbide threads or fibers. The flow directing guide rails 80 are arranged to cause the fuel to rotate as it moves from the manifold 50 toward the burner section 64 of the rocket housing 8. The flow directing guide rails may be provided throughout the fuel passage or in segments thereof. In most preferred embodiments of the invention, the guide rails 80 are located at least in the fuel passage 56 within the area of the combustion chamber 64.

Fuel injection openings 68, preferably in the form of circular holes in the wall S2 within the area of the combustion chamber 64 are arranged in the wall 52 in such a way that they continue or promote the vortex movement of the hot gaseous fuel represented by the arrows 78 as it passes from the fuel passage 56 into the combustion chamber 56. . Any desired number of fuel injection openings 68 may be used in accordance with the present invention. In most embodiments, the injection rate of the hot gaseous fuel into the combustion chamber exceeds about 300 m / s (1000 feet per second) at about 1500 kp / mm 2 (1000 p.s.i.) pressure drop. A plurality of fuel injection holes 68 are provided in the burner liner 52 for injecting fuel into the combustion chamber in a direction which supports the peripheral movement of the fuel. In preferred embodiments, where the burner liner is substantially circular in cross section, the fuel injection holes in the burner liner are oriented in the burner liner at an angle less than 90 ° to a perpendicular line to the tangent of the burner liner. In most embodiments, this angle is about 300 to about 600. Any angle less than 900, which is sufficient to support the peripheral movement of the fuel as it enters from the fuel passage into the combustion chamber can be used.

The propulsion system of the present invention also includes means for introducing oxidant, such as liquid oxygen, from the oxidizer line 72, which is provided with oxidizer from the oxidizer tanks 86 shown in Figures 1 and 2.

The oxidizer is introduced into the combustion chamber through oxidizer injection openings 66 as shown in Figs. 4 and 5. The oxidizer line 72 is centrally located in the combustion chamber and has a plurality of oxidizer injection holes or openings 66 for introducing oxidizer to the center of the combustion chamber. The oxidizer is preferably injected into the combustion chamber in a direction which supports the peripheral movement of the oxidizer as it enters the combustion chamber from the oxidizer line 72, so that fuel and oxidant are mixed in the chamber before and during combustion. In preferred embodiments, the oxidizer insertion device into the combustion chamber is substantially circular in cross section, and the oxidizer injection holes or openings 66 are oriented at an angle less than 900 ° to a perpendicular line to the tangent of the oxidizer insertion device into the chamber. However, in preferred embodiments, the angle of the oxidizer injection ports is about 300 to about 600. Any angle sufficient to support the peripheral movement of the oxidizer as it enters the combustion chamber may be used. In certain preferred embodiments of the present invention, the propulsion system comprises a plurality of fuel injection openings 66 and oxidizer injection openings 68 arranged in a zigzag to promote overlap of fuel and oxidant in the combustion chamber and further promote mixing of oxidant and fuel before and during fuel. In this way, fuel injected through the injection openings 68 and oxidizer injected through the injection openings 66 so that mutual penetration occurs between the fuels and oxidizer before and during combustion. In this way, a highly developed state of mixing and rapid combustion is achieved when the fuel mixture is ignited.

The fuel injection line 72 comprises a porous sleeve, which is preferably made of woven silicon carbide fibers. Hydrogen 70 is introduced through hydrogen injection lines 74. The porous hydrogen injection line 74 also provides the perspiration cooling of the injection line 74 as discussed above for the burner liner 52 and nozzle neck 52. The hydrogen 70 introduced through the hydrogen injection ports 76 is also mixed with the oxidizer to initiate combustion and cooling line.

In certain embodiments of the present invention, the peripheral movement of the fuel from the fuel injection hole in the burner liner is substantially clockwise, and the peripheral movement of the oxidizer through the oxidizer injection holes into the combustion chamber is substantially counterclockwise. In other embodiments, the peripheral movement of the fuel through the fuel injection holes in the burner liner is substantially counterclockwise, and the peripheral movement of the oxidizer through the oxidizer injection holes in the oxidizer insertion device is substantially clockwise.

According to the present invention, a propulsion system is obtained which utilizes a hydrocarbon fuel, wherein means are provided for providing high heat flow to the burner liner and neck liner, which heat is obtained from combustion within the rocket casing or rocket chamber and nozzle neck, the thermal temperature of the liner exceeding due to combustion in the rocket casing, as well as means for cooling the burner liner and the nozzle neck liner by endothermic pyrolysis of hydrocarbon in the fuel passage. Thus, in addition to convection and film cooling by the fuel, as discussed above, decomposition, cracking and / or dehydration of hydrocarbon by endothermic pyrolysis occurs in the fuel passage 56 prior to combustion. Using ethylene (CZH4) as an example, the following reaction occurs in the fuel passage in the presence of heat and in the absence of catalyst: 464 719 16 2C2H4 -> C4Hs HEAT C4H8 ___) H2 + C4H6 HEAT H2 + C4H5 V2M2H2H2H2 The above endothermic reaction gives 0.9 kg (2 pounds) of hydrogen and 12 kg (26 pounds) of acetylene gas (CZHZ) of 12.7 kg (28 pounds) of ethylene (C2H4) - The heat absorbed during the reaction is 1340 kcal / kg (2413 BTU per pound) ethylene.

Due to its extremely high combustion temperature, acetylene is only inferior to hydrogen in terms of maximum rocket traction and specific impulse. Thus, according to the present invention, ethylene is the most preferred resulting product from endothermic pyrolysis of the hydrocarbons.

Thus, when the hydrocarbon fuel is ethylene, the products of endothermic pyrolysis include acetylene and hydrogen. When the hydrocarbon fuel is ethane, the products of endothermic pyrolysis include methane, ethylene and hydrogen as well as trace amounts of various other hydrocarbons. When the hydrocarbon is propylene, the products of endothermic pyrolysis include methane, ethane, ethylene, acetylene and hydrogen and trace amounts of other hydrocarbon products. When the hydrocarbon fuel is propane, the products of endothermic pyrolysis include methane, ethane, ethylene, acetylene, propylene and hydrogen and trace amounts of various other hydrocarbon products. Of course, it is within the scope of the present invention to use different mixtures of the above hydrocarbons. In any case, the hydrocarbon fuels used in the present invention are those in which endothermic pyrolysis produces a fuel product with higher combustion rates, a fuel product with higher combustion temperatures and / or a fuel product with a lower molecular weight than the hydrocarbon fuel from which it is derived. As stated above, the fuel products are formed in the fuel passage 56 in the presence of the heat generated in the fuel chamber and the nozzle and which is conducted through the wall 52.

Catalyst can be used in conjunction with the hydrocarbon fuel to accelerate the change in the composition of the hydrocarbon fuel, thus providing additional endothermic cooling of the wall 52 and in some cases, such as with acetylene, a superior rocket fuel product is obtained from the endothermic pyrolysis. For example, the catalyst device for the endothermic pyrolysis may be a catalyst bed in the fuel passage 56, as shown in Fig. 5, where the catalyst 84 is shown in the form of balls located in the passage. The catalyst may also be coated on a wall of the fuel passage 56, such as the wall 54 and / or the wall 52. The catalyst u: 464 719 g 17 may also be included in the hydrocarbon fuel so that when the fuel is driven through the fuel passage 56 the catalyst in the fuel accelerates the endothermic pyrolysis of hydrocarbon - the fuel. Any catalyst of the type well known in the art of endothermic pyrolysis and cracking can be used in the present invention, including such catalysts as platinum and palladium.

In the above process, energy is taken from the burner and nozzle neck through the wall 52 to the fuel in the fuel passage 56, then back to the combustion chamber through the fuel injection ports 68. There is no net loss or gain of heat energy, but the wall 52 is cooled by the endothermic and reactive thermal reaction. combustion temperatures and combustion products are obtained, whereby higher rocket nozzle velocity and higher traction are obtained.

Thus, the present invention provides a method of improving the fuel system of a propulsion system having a rocket housing having a combustion chamber having a burner liner and an exhaust gas neck having a neck liner, a fuel passage being adjacent to and surrounding the burner liner and neck liner. The hydrocarbon fuel is led through the fuel passage. Heat from the combustion of fuel in the combustion chamber is emitted to the fuel passage by radiation through the burner lining and the neck lining. The hydrocarbon fuel is heated at a sufficient temperature to cause endothermic pyrolysis of the hydrocarbon in the fuel passage, whereby heat removed from the combustion chamber and the neck liner reduces the temperature in the combustion chamber and the burner liner and neck liner so that the thermal constraints of the burner liner neck are exceeded.

In addition to the propellant flow rate, which determines the combustion pressure (rocket neck throttled at speed of sound) and the mixing ratio, which determines the combustion temperature, the ratio of hydrogen to hydrocarbon fuel can be varied to control the temperature of the structure (except rocket traction).

Excess hydrogen increases the film cooling, reduces the combustion temperature (fuel-rich), reduces radiation from carbon compounds (hydrocarbon fuel) and promotes the formation of water with a relatively low molecular weight. The ratio of hydrogen to hydrocarbon can be adjusted by those skilled in the art to achieve the most desirable combination of the foregoing variables, but in most cases a mass ratio of 50% by weight hydrogen and 50% by weight hydrocarbon has generally been found to be desirable.

The present invention proposes a systematic use of hydrocarbon and hydrogen fuel, and in doing so there is no limitation on the devices by which hydrocarbon is decomposed or converted to any other form of hydrocarbon. For example, the conversion of ethylene or ethane to acetylene is well known in the art. However, this conversion typically involves 464,719 ”in difficulty in achieving maximum conversion yield (about 70% yield per unit weight of acetylene from ethylene) within a short period of time (less than 5 milliseconds) at elevated temperatures. Furthermore, the conversion process, whether pyrolytic and / or catalytic, can form undesirable carbon products, such as carbon-carbon molecules (coke) or metal carbides (of the catalyst). A particularly unfavorable reaction is acetylene explosion due to carbon-carbon polymerization.

According to one aspect of the present invention, the hydrocarbons are converted to more favorable products, i.e. fuel products with higher combustion rates, higher combustion temperatures and / or lower molecular weights, by thermal conversion at about 925 ° C (1700 ° F) to about 1200 ° C (2200 ° F), using the hydrogen fuel already available in the fuel mixture obtained from the manifold 50, to accelerate a molecular reaction, where a single hydrogen molecule is dissociated from the hydrocarbon and combined with two hydrogen atoms from the hydrocarbon to form a more desirable product, namely acetylene. At about 925 ° C (1700 ° F), for example, ethylene in the presence of hydrogen forms acetylene.

The activation energy is reduced and the conversion speed is accelerated. In addition, the excess hydrogen buffers the formation of carbon-carbon molecules including carbon polymerization because hydrogen molecules fit within the free average path length of free carbon atoms and therefore block one carbon atom from another carbon atom. Of course, additional catalyst devices can also be used with the means for cooling the burner liner and the neck liner by endothermic pyrolysis of hydrocarbon in the presence of hydrogen in the fuel passage, the hydrogen in the fuel accelerating the rate of the endothermic pyrolysis. When the hydrocarbon fuel is a hydrocarbon, the endothermic pyrolysis of which results in the formation of coke and carbon polymers, the hydrogen in the fuel also suppresses the formation of coke and the polymerization of carbon, thereby preventing explosion in the fuel passage. Conventional catalyst devices as above can also be used in this propulsion system. Although not intended to limit endothermic pyrolysis to any particular temperature, according to the present invention, the hydrocarbon fuel and hydrogen are generally heated in the fuel passage at a temperature of about 925 ° C (1700 ° F) to about 1200 ° C (2000 ° F). ). Although the hydrocarbon in the fuel tank is a gas or a liquid or a mixture thereof, and although the hydrogen in the fuel tank is a liquid or semi-solid, the hydrocarbon and hydrogen in the fuel passage 56 are generally in a gaseous state due to the high temperature in the passage.

Although not shown, those skilled in the art can easily insert means for igniting the fuel in the presence of oxidizer in the combustion chamber. The igniter may be, for example, an electrical system which provides an electric arc within the range of the fuel injection system, for example the hydrogen openings 76, the fuel injection openings 68 and the oxidizer injection openings 66.

The properties and characteristics described above individually or in combination provide improved propulsion systems and methods which utilize the hydrocarbon fuel or hydrocarbon fuels in combination with the hydrogen fuel in the presence of oxidizer. The above-mentioned systems also enable the use of improved aircraft with propulsion systems, which utilize a twin fuel system and a method of storing fuel in a vehicle whose propulsion system uses a twin fuel system.

Claims (6)

464 719 2 ° Patent claim. Propulsion system comprising: a) a rocket casing (8) comprising a combustion chamber (64) having a burner liner (52) and a nozzle neck chamber (62) having a nozzle neck liner f (52), b) a fuel passage (56) in the vicinity of and surrounding the burner liner and the nozzle neck liner, which form the inner wall (52) of the fuel passage, the outer wall (54) of the fuel passage being separated from the inner wall to form the passage, c) means for providing fuel flow in the fuel passage, and characterized by: d ( ) in the burner liner (52), which injects fuel into the combustion chamber (64) in a direction which promotes peripheral movement of the fuel, and e) means (70) for introducing oxidizer through oxidizer injection holes (66) to the center of the combustion chamber so that oxidizer is injected in the combustion chamber in a direction which promotes peripheral movement of the oxidizer so that the fuel and oxidizer are mixed in the chamber before and during combustion. Propulsion system according to claim 1, characterized in that the fuel injection holes (68) and the oxidizer injection holes (66) are zigzagged or offset arranged to promote fuel and oxidant overlap and mixing before and during combustion. Propulsion system according to claim 1, characterized in that the burner liner (52) is substantially circular and the fuel injection holes (68) in the burner liner are oriented at an angle of less than 900 to a line perpendicular to the tangent of the burner liner. Propulsion system according to claim 1, characterized in that said means (70) for introducing oxidizer is substantially circular in cross-section, and that the oxidizer injection holes (66) are oriented at an angle of less than 900 to a perpendicular line to the tangent of said means (70) for introducing oxidizer. Propulsion system according to claim 1, characterized in that the peripheral movement of the fuel from the fuel injection holes (68) in the burner liner a. §X (52) is substantially clockwise, and the peripheral movement of the oxidizer through the oxidizer injection holes (66) into the combustion chamber is substantially counterclockwise . 21 6 464 719 Propulsion system according to claim 1, characterized in that the peripheral movement of the fuel through the fuel injection port (68) in the combustion chamber is substantially counterclockwise, and the peripheral movement of the oxidizer through the oxidizer injection means is substantially inward to the oxidizer hole.