FR2610044A1 - PROPULSION SYSTEM COMPRISING AN IMPROVED TRIM OF THE COMBUSTION CHAMBER AND METHOD OF MANUFACTURING SUCH MANUFACTURE - Google Patents

Abstract

AN INTERNAL WALL WOVEN OF SILICON CARBIDE FIBER CONSTITUTES A TRIM 52 OF THE COMBUSTION CHAMBER 64 AND A TRIM FOR THE COL OF THE TUBE 62, AND AN OUTER WALL WOVEN IN SILICON CARBIDE FIBERS IS SPACED FROM THE INTERIOR WALL CHANNEL 56 FOR FUEL. THE SILICON CARBIDE FIBERS ARE WOVEN SO AS TO EXTRACT BY CONDUCTION THE HEAT FROM THE COMBUSTION CHAMBER AND FROM THE TUBE NECK TO TRANSMIT IT TO THE FUEL CHANNEL AND THEREFORE PROVIDE HEAT FOR THE ENDOTHERMAL PYROLYSIS OF THE FUEL SYSTEM , WHICH INCLUDES A ROCKET ENCLOSURE CONTAINING A COMBUSTION CHAMBER, FUEL INJECTORS, A TUBE NECK AND A PIPE. THE FUEL CHANNEL IS CONTIGUOUS, BY SURROUNDING THEM, TO THE COMBUSTION CHAMBER AND TO THE TUYERE THROUGH, AND THE FUEL IS SUBJECTED TO ENDOTHERMAL PYROLYSIS IN THE CHANNEL. APPLICATION TO ROCKET ENGINES, STATOREACTORS, ETC.

Description

The present invention relates to a system and a

  propulsion process for aircraft and, more particularly,

  a propulsion system and process for an airplane

  equipped with rocket motors, ramjets or

  supersonic combustion reactors. Liquid hydrogen, a liquid hydrocarbon and liquid oxygen are stored and used in a systematic manner to meet the conditions corresponding to high propulsion thrust, low vehicle weight and low vehicle drag. an aircraft and system realization

relatively simple propulsion.

  Flight to a single-stage orbit in high accelerations can be considered as the combination of a high available net thrust,

  vehicle weight and low air drag.

  dynamic. The weight of the fuels decreases while the

  flight progresses and with a positive net thrust, the vehicle

  will reach the orbital insertion rate at altitudes

  orbital studies to the extent that there are reservations in

  fuels. Therefore, a key factor in the propulsion

  An aircraft is the amount of fuel in relation to the dimensions of the vehicle. A second key factor in a one-step flight to orbit is a fixed weight

  low value for the vehicle and the propulsion system.

  - 2 - sion because the thrust must take account of this weight during the entire ascent until the setting into orbit. A third factor in this type of theft is the shape of the vehicle that

  must achieve a favorable relationship between

  sional and drag. The final key factor is a high thrust on the part of the propulsion system, that is to say, the rocket engines of the vehicle. In view of the above, it is easy to see that it is desirable to increase the

  amount of fuel in relation to the dimensions of the vehicle

  to reduce the fixed weight of the vehicle and the propulsion system, thus reducing the thrust required to propel the vehicle and the propulsion system to the orbit, to increase the aerodynamic shape of the vehicle. aircraft, and increase the thrust that can be achieved with the system

propulsion.

  Many liquid propulsion systems, including liquid multi-fuel propulsion systems, are known in the art. However, it is difficult to store the fuels in the vehicle to achieve maximum utilization of the storage space in this vehicle while reducing the weight of the tanks and conferring the vehicle a low value aerodynamic drag. In order to achieve the above objectives for liquid fuel propulsion systems, it has generally been necessary to

  vacuum type storage in vehicles for the storage of

  coking of liquid fuels. Containers and systems

  vacuum storage are difficult to maintain and increase

  significantly lower the weight of the vehicle. Therefore, it is desirable to avoid the presence of vacuum type storage tanks and other large systems

  for the storage of liquid fuels on board

  naves. It has been widely used as fuel in

  aircraft propulsion systems for hydrocarbon mixtures

26 10 0 4 4

  - 3 - bure-liquid oxygen. It is known that such hydrocarbon fuels burn in the combustion chambers

  at temperatures exceeding 2700 C and producing a

  high. However, propulsion systems for newer aircraft require a higher thrust

  the one most often obtained with the hydro-

  carbide-oxygen liquid and, in general, has been developed

  more complex fuels in propulsion systems.

  more complicated to provide the higher thrust. Most of these complex fuels can not be used in conventional propulsion systems for various

  reasons, among other things, because of the

  extremely high levels, and propulsion systems and storage systems need to be

  complex to take into account the combustion of such

  patible. The complex propulsion systems that allow

  to overcome the above difficulties,

  the weight of the vehicle. Given the above parameters, it would be advantageous to use

  less complex propulsion based on hydrocarbon mixtures

  bure-oxygen liquid to avoid inconvenience

to be exposed.

  It is known that certain chemicals, such as high molecular weight hydrocarbons, react endothermically to produce reaction products that can be used as fuel. We developed in art

  systems that convert certain subsystems

  chemical reactions by endothermic reactions in

  that can be burned in a combustion chamber.

  However, most systems of the prior art are disadvantageous because the process of the endothermic reaction gives certain products which can accumulate in considerable amount in the combustion chamber. Therefore, it is desirable to provide a combustion system that uses simple hydrocarbons, such as ethylene, ethane, propylene and propane, to constitute the chemical that reacts endothermally so as to give first and foremost reaction products that can be used as advanced fuels in aircraft propulsion systems. In the prior art propulsion systems which use fuels producing high temperatures in the propulsion system, for example in the combustion chamber, it has proved difficult to find materials capable of withstanding the high temperatures. that is, 2700 ° and over Many types of fillings for the combustion chambers and nozzle necks of rocket casings have been proposed, however, they are complex and expensive and have a limited service life. require excessive amounts of refrigerant to be circulated to prevent weakening of structures and / or melting of the packing material, therefore, it is desirable to provide improved combustion chamber and collar designs of nozzle and materials in propulsion systems that avoid

the previous disadvantages.

  Accordingly, the subject of the present invention is

  main aircraft and a propulsion system perfected

  of the type using multiple fuels.

  The present invention relates to a propulsion system and method using a fuel choice systematically to obtain a flight in one step

  to orbit while taking into consideration

  implementation of the aircraft and the propulsion system.

if we.

  The present invention further relates to a fuel storage system and method for storing multiple fuels in a single, lightweight container, and providing maximum fuel reserves in a

  an aircraft of minimum size and weight.

-5-

  Another object of the present invention is a provision

  and a fuel storage method that provides maximum fuel reserves while imparting optimum shape to the aircraft in terms of climbing power and drag. Still another object of the present invention is a propulsion system and method utilizing combustion chamber liners and tuyere neck liners for retention of high combustion pressures and

  can withstand high combustion temperatures.

  Another object of the present invention is a system

  a propulsion process using a mixture of hy-

  hydrocarbon and hydrogen with an oxidant based on liquid oxygen for high combustion temperatures from

  of low molecular weight fuels.

  The present invention further relates to a propulsion system and method for cooling a lining of the combustion chamber and the neck of

nozzle of a rocket engine.

  The present invention also relates to a system

  and an improved method of propulsion in which the combustion chamber liner and the throat nozzle of a rocket engine are cooled by the

  endothermic decomposition of the fuel.

  The present invention further relates to a combustion chamber and nozzle neck of improved design that increases the heat transfer and residence time for the endothermic cooling of the lining of the combustion chamber and that of the

neck of the nozzle.

  The present invention also relates to a system

  advanced fuel injection system and system

  perfected oxygen injection to proceed with the ejection

  fuel and liquid oxygen in the chamber

  of combustion of a propulsion system.

  -6- We reach these objects as well as other objects

  in the present invention by choosing a specific class

  of fuels used in conjunction with a

  advanced propulsion system and an improved design

of the aircraft.

  The advanced propulsion system of the pre-

  this invention uses a hydrocarbon fuel and employs a rocket shell having a combustion chamber liner and a nozzle neck; a channel for the adjoining hydrocarbon, surrounding them, to the combustion chamber liner and the nozzle neck liner; means for supplying a fuel stream to the channel; means for providing a high heat flow to the combustion chamber liner and nozzle neck liner from combustion in the casing

  of the rocket, from which it follows that the temperature of the

  temperatures exceed their thermal limits; and means for cooling the combustion chamber liner and the neck of the nozzle by endothermic pyrolysis of the hydrocarbon in the channel. In one of the aspects of the

  the present invention, the endothermic pyrolysis of

  in the channel is carried out in the presence of hydrogen, and the hydrogen of the fuel accelerates the rate of pyrolysis.

  endothermic. In another aspect of the present invention

  the endothermic pyrolysis of the hydrocarbon takes place in the channel in the presence of a catalyst means which

  accelerates the speed of this pyrolysis. Endothelial pyrolysis

  of the class of hydrocarbons used in the pre-

  invention, in the presence of hydrogen and / or

  of a combustible product having higher combustion rates, a combustible product having higher combustion temperatures and / or a product

  fuel having a lower molecular weight.

  The aircraft of the present invention comprises a propulsion system which uses a dual fuel system, wherein at least one of the fuel is a cryogenic fuel, a non-evacuated inner tank contains the cryogenic fuel, and an outer reservoir, which is not under vacuum and surrounds the inner reservoir, contains a second fuel, the second fuel having a low freezing point and a

  high boiling point, which acts by isolating for the com-

  cryogenic fuel contained in the inner tank. The second fuel in the outer tank may also be a cryogenic fuel. The fuel tank or container system comprises a pressurized cylindrical inner tank or container for holding or storing liquid cryogenic fuel, for example

  hydrogen, and comprises a rigid insulated wall. In total-

  In the inner tank for holding and storing liquid cryogenic fuel, an outer tank contains the second fuel which is generally a liquid and / or gaseous hydrocarbon. By surrounding the inner reservoir, for example containing liquid hydrogen, with a pre-cooled liquid and / or gaseous hydrocarbon, excessive boiling of liquid hydrogen at high altitudes is avoided. The liquid and / or gaseous hydrocarbon layer isolates the liquid hydrogen, thereby avoiding conventional cryogenic vacuum containers for isolation. The outer tank wall that contains

  the hydrocarbon is the coating of the aircraft.

  Thus, the outer wall of the outer tank matches the aerodynamic shape of the aircraft. The pre-cooled hydrocarbon is substantially non-pressurized and therefore fills the entire space between the liquid hydrogen reservoir and the aircraft liner. This design and method of storing the fuel in the aircraft makes it possible to give it a shape resulting in an ascending power and a drag

  optimum without reducing fuel reserves.

  -8 - Storage means for a liquid oxidant

  are provided, and conventional means used to

  hold the liquid oxidizer in a sub-cooled state to reduce boiling. The liquid oxidant storage means may be of any suitable shape to fill one or more non-pressurized containers which conform to the shape of the aircraft and the improved system and

  fuel storage of the present invention.

  According to the present invention,

  a propulsion system having a rocket envelope con-

  holding a combustion chamber, fuel injectors

  tible, a nozzle neck and a nozzle; a channel for

  fuel in which the fuel is subjected to pyrolysis

  endothermic lysis, this canal being contiguous, surrounded by

  rant, to the combustion chamber and the neck of the nozzle; and means for supplying a fuel stream into the channel, the improvement comprises an inner wall of woven silicon carbide fibers which constitutes a lining of the combustion chamber and a neck liner of the nozzle and a fiber outer wall woven silicon carbide spaced from the inner wall to form said fuel channel. The silicon carbide fibers are woven, preferably continuously, and conduct the heat between the combustion chamber and the

  from the nozzle to the fuel channel, hence the

  port of heat for the endothermic pyrolysis of

  tible. In preferred embodiments, the inner wall of woven silicon carbide fibers has a certain porosity for selectively controlling the diffusion of hydrogen between the channel, via the wall, and the combustion chamber. the neck of the nozzle. This improvement of the propulsion system, that is to say in rocket engines in which the combustion chamber consists of continuously wound silicon carbide fibers, makes it possible for the electric motors of 261004t and

  operate at very high temperatures with combus-

  Such low molecular weight, i.e., for example, fuels such as low molecular weight hydrocarbons and hydrogen in the presence of the oxidant, namely oxygen. The products of combustion of these fuels have a relatively low molecular weight. As they are used here, hydrocarbons to

  low molecular weight are unsaturated or are hydro-

  saturated carbides having less than 4 carbon atoms.

  Hydrogen from liquid hydrogen in the hydrocarbon mixture in addition to hydrogen due to the endothermic decomposition of the hydrocarbon having the fuel channel promotes the formation of a large fraction of mass water vapor relatively low molecular weight in the products of combustion. The high combustion temperature in the combustion chamber and the high temperature in the neck of the nozzle are taken into account by the combined use of the continuously wound silicon carbide fibers and the fluid film cooling due to the Hydrogen present in the channel in which the hydrogen diffuses between the channel, through the firebox lining and the nozzle neck liner, and the combustion chamber and the nozzle neck chamber , respectively. The net result is a combustion temperature and a ratio of molecular weights that cause high sonic velocity in

  the neck of the engine nozzle rocket, and after relaxation super-

  sonic, a high speed at the engine-rocket exhaust.

  According to the present invention, the construction of wound silicon carbide fibers of the combustion chamber lining and the nozzle neck liner provides high peripheral resistance, and the

  layers of fiber provide cooling by transpiration

  ration thanks to the controlled porosity of the fibers, which

  allows the migration or diffusion under pressure of the hydro-

- 10 -

  gene, and not the hydrocarbon, into the combustion chamber and the chamber of the neck of the nozzle through their walls. Thus, a fluid film cooling is obtained on the hot side of the wall while simultaneously the hydrogen addition decreases the molecular weight of the exhaust gas, resulting in a higher velocity at the neck. In addition, the hydrogen film on the combustion chamber side and nozzle throat of the chamber liner and the neck liner reduce friction of the fluid on the wall

the neck of the nozzle.

  In another aspect of the present invention, the propulsion system comprises a rocket casing having a combustion chamber with a liner and a chamber

  nozzle neck with a liner; a fuel channel

  adjacent to the firebox lining and the nozzle throat

  which form the inner wall of the canal, the outer wall

  of this channel being spaced from the inner wall for

  form the channel; means for providing a current of communication

  bustible to the canal; and a multitude of flow guide vanes which are arranged in the channel to direct the flow.

  circumferentially combustible. Thus, the fuel

  A longer path in the canal increases the residence time and facilitates the transfer of heat to the fuel in the canal. Thanks to the multitude of fins, there is therefore a longer residence time

  in the fuel channel, and this allows for a trans-

  higher heat density and greater time for endothermic pyrolysis or hydrocarbon decomposition

serving as fuel.

  In another improvement, a plurality of fuel injection holes in the combustion chamber liner inject fuel into the chamber in a direction that promotes circumferential fuel movement in the chamber. An oxidant

- he -

  is also introduced through

  and directed towards the center of the combustion chamber so as to inject the oxidant into the chamber in a direction which promotes the circumferential movement of this oxidant. In this way, the fuel and the oxidant are mixed in the chamber before combustion and during

of it.

  Thanks to the present invention, the hydrocarbon

  fuel is used to generate

  goods having at least one of the following properties: (1)

  higher burning rate, (2)

  bustion higher and (3) lower molecular weight

  before burning in the room. For example, acethyl

  is a preferred fuel product that has a

  higher burning rate, a com-

  higher boiling point and lower molecular weight than the hydrocarbon from which it is produced. In one embodiment of the present invention, ethylene is thermally converted to acetylene and hydrogen using hydrogen in the fuel channel as a catalyst and / or employing an additional catalyst,

  while excess hydrogen from the fuel mixture

  tible also acts to buffer and prevent the polymerization of carbon which results in coke or in some cases

  by the detonation of acetylene.

  Thanks to the present invention, it is possible

  to increase the consumption of liquid oxygen, hydrocar-

  liquid and / or gaseous and liquid hydrogen, in this

  order. This is important because for any speed of escape

  of a rocket, the thrust is obtained by adding the

  mass of fuels so as to increase the moment.

  The liquid oxygen, followed by the liquid hydrocarbon, has a high density in the liquid state, and therefore a high fuel fraction relative to the

  the aircraft and the fixed weight of the latter.

2610 0 4 4

- 12 -

  The propulsion systems and methods of the present invention can be used in any aircraft whose engine (s) uses a rocket envelope; for example, the propulsion systems and methods of the present invention may constitute a part of a rocket engine, part of a ramjet or part of

  a supersonic combustion ramjet.

  The following description refers to the figures

  Figs. 1, an isometric illustration, partly broken away, of a rocket-driven aircraft which illustrates the fuel containers of the present invention; Figure 2 is a sectional view taken along the lines 2-2 of Figure 1, illustrating the fuel and oxidant tanks of the aircraft;

  FIG. 3 is a sectional view of a segment of the reservoirs

  see fuel, illustrating the construction of their walls; 4, a sectional view of the upper part of a rocket engine and illustrating the fuel channel, the nozzle neck lining, the lining of the chamber of

  combustion and the oxidant injector according to the present invention.

  tion; FIG. 5 is a sectional view of the rocket engine taken along the lines 5-5 of FIG. 4, and illustrating in FIG.

  in addition to the fuel flow in the canal and the

  flow control letters of the present invention; Figure 6, a sectional view of the filling of the

  combustion chamber of the present invention.

  Figure 1 shows a typical aircraft having a cabin and / or load 24 in its anterior section, three engines 8 and a rear section 22

  in the rear of the aircraft and the improved system

  fuel storage system of the present invention which includes a liquid hydrogen reservoir 4 surrounded by

- 13 -

  by a hydrocarbon reservoir as well as by reservoirs

  see 6 liquid oxidant located in the middle of the rear section of the aircraft. The advanced propulsion systems and methods of the present invention can be used in any aircraft embodiment. In connection with FIGS. 1 and 2, the liquid and / or pasty hydrogen is contained and pressurized in the

  tank 4 of the aircraft driven by rocket or airplane 10.

  Pasty hydrogen is a combination of liquid hydrogen

  and solid hydrogen. The hydrocarbon reservoir 20

  the tank 4. As used here, the tank

  see 4 is a non-evacuated inner container which contains a cryogenic fuel, such as hydrogen, normally requiring an isolation chamber or container

  under vacuum, and the hydrocarbon reservoir 20 is a reservoir.

  see outside, not subject to vacuum, which surrounds the inner tank to enclose a second fuel, for example a fuel having a low freezing point and a

  boiling point which acts by isolating for the combustion

  cryogenic solution contained in the inner tank 4.

  Examples of fuels having a low freezing point include those having a point

  freezing rate between about - 190 ° C and about -

  C and a high boiling point, for example between about -120 ° C and about -37 ° C. Hydrocarbons, whether in the liquid state or in the gaseous state, and whose freezing and boiling points fall into the

  previous ranges, provide the necessary thermal barrier

  to avoid the boiling of hydrogen. Isolation of the wall separating the liquid hydrogen reservoir 4 from the hydrocarbon reservoir is sufficient to prevent freezing of the hydrocarbon by the liquid hydrogen during the normal duration of a flight. At an absolute pressure in the reservoir of 0.14 MPa, the liquid hydrogen has a temperature of -250 C absorbs 5.3 kJ per 450 g, hydrogen

- 14 -

  pasty at a temperature of -260 C absorbs 26 kJ per 450 g before boiling occurs. In addition to a low point

  freezing, the hydrocarbon must have a boiling point

  high pressure (low vapor pressure) so as not to

  no pressurization at high altitudes. According to the present invention, pressurization at less than about

  0.014 MPa absolute (pressure above the ambient pressure)

  biante) is acceptable. The low vapor pressure of the hydrocarbon allows the hydrocarbon reservoir to have the desired shape. Therefore, the reservoir 20 can be shaped to provide the aircraft with reduced aerodynamic drag and lower heating by

friction in the atmosphere.

  The hydrocarbon used in accordance with the foregoing parameters to provide the necessary isolation for liquid hydrogen include ethylene, ethane, propylene, propane, and mixtures thereof. As

  these fuels are also cryogenic fuels

  However, the fuel of the outer tank 20 may also be a cryogenic fuel. The following table gives examples of typical hydrocarbons that can be

  use in the tank 20 according to the present invention.

BOARD

HYDROCARBONS

  Point of Density at the boiling point liquid freezing at -140C (C) (C) g / cm3) Ethylene-104C -169C 0.61 C2H4 Ethane-93C-172C 0.61C2H6 Propylene -48C -185C 0, 64 C3H6 Propane -40 C -190 C 0.62 C3H8

- 15 -

  The fuels indicated in the previous table

  The tooth can be subcooled if necessary, to reduce their vapor pressure and prevent boiling at high altitudes. For example, the table fuels can be subcooled to -140 C. The density in the liquid state at -140 ° C is indicated for each

hydrocarbons from the table.

  The inner tank 4 which contains the liquid hydrogen is generally cylindrical as illustrated in FIGS. 1 and 2. In the preferred embodiments,

  the tank 4 is an elongated cylinder which extends practically

  between the cabin and the load section 24 over the entire length of the aircraft to the section containing the rocket. For example, the tank 4 is located in the center of the aircraft as shown in FIG.

  approximately between line 26 and line 28 of the

  As shown in FIG. 1, a pipe 16 of liquid hydrogen is connected to a conventional pump 18.

  to supply liquid hydrogen to pipelines and

  appropriate authorities to distribute it to the

  channels of the propulsion systems of the present invention.

  Although there is shown in Figure 1 a single line of liquid hydrogen and a single pump, the skilled person will provide a multitude of pipes and fuel pumps, as well as ancillary equipment and

  orders by allowing distribution.

  The liquid hydrogen reservoir 4 can be

  made of any insulating material, the thickness of which

  It will prevent liquid hydrogen freezing of the hydrocarbon contained in the tank 20 during the flight. As shown in FIG.

  light construction typical of the wall of the reservoir of

  liquid drogen comprises walls 38 made of graphite-epoxy resin, preferably of reinforced fibers, for example from

  graphite fibers, which are separated by an insulating material

- 16 -

  40. For example, the insulating material may be

  killed about 2.5 cm of liquid polyurethane foam injected into the walls 38 and cured therebetween to form the polyurethane insulation 40. Other typical insulating materials that can withstand low temperatures include, for example, packed silica. The insulating material 40 prevents the freezing of the liquid hydrocarbon

* in the tank 20 by the liquid hydrogen of the tank 4.

  The inner surface of the reservoir 4 is preferably coated with a material preventing the diffusion of hydrogen

  through the walls 38. For example, a metal foil

  A nickel-iron alloy containing between about 40% and about 50% nickel and having a low coefficient of thermal expansion is suitable for the coating 42 of the inner wall 38 of the tank 4. A foil can

  to be used for this purpose is sold under the brand Invar.

  According to the present invention, the outer wall of the outer tank, that is to say the outer wall of the tank 20 is the coating 30 of the aircraft 10. Thus, the

  outer wall of the tank 20 matches the aerodynamic shape

  10. The coating 30 of the aircraft 10 preferably has a thickness of about 2.5 cm. Typical coating materials are well known in the art. For example, materials for high temperatures (about

  870 "C) such as titanium aluminite or certain superalloys

  Nickel-based bindings, such as the Rene 4, can be used to form the vehicle's coating. The outer covering 30 is bonded to the high temperature resistant insulating material 36, such as thermal insulation having the trade name Min-K reinforced with a fibrous medium and very fine particulates heat resistant and having a micro structure. porous with very low thermal conductivity and low

  thermal diffusivity. It is also possible to use

261004t

- 17 -

  crushed silica, for example compacted silica

  containing a fibrous medium and a very fine material parti-

  heat-resistant ring having a micro-structure

  porous with low thermal conductivity and low thermal diffusivity to form the insulation material

  36. The inner surface of the outer wall of the tank

  see 20 is also covered with a suitable material to reduce radiation. For example, an iron-nickel alloy containing between about 40% and about 50% nickel in the form of a foil can be used to provide the radiation barrier 34 and coating 34 of the wall of the tank 20. 34 tinsel laminated with a

  high temperature insulation 36 also reduces the radiation

  ments. The metal foil 34 also serves to reduce

  the gradients of thermal expansion in the tank.

  In order to avoid the freezing of moisture on the outer skin 30 of the aircraft 10, a purge gas is used in the spaces 32 of the structure 31 of the grid. For example, hot nitrogen gas may be supplied as purge gas to the spaces 32 of the grid 31.

  in order to heat the outer coating 30.

  According to the present invention there is provided a fuel storage method in an aircraft comprising a propulsion system which uses a dual fuel system in which one of the fuels is a cryogenic fuel which normally requires a cryogenic insulation

  under vacuum, including the placement of the cryogenic fuel

  normally requiring a cryogenic vessel under vacuum in a non-evacuated inner tank, and placing a second fuel having a low freezing point and a high boiling point in the area surrounding the inner tank, d It follows that the

  second fuel acts as an insulator for the fuel

  the cryogenic solution contained in the inner tank.

  In FIG. 1, a pipe 17 is connected to a

261004L

- 18 -

  hydrocarbon pump to supply it to the appropriate lines and manifolds according to the present invention. A line 12 supplies liquid oxygen to the appropriate lines and manifolds by means of a pump 14. Any conventional fuel delivery system, including

  fuel lines, as well as controls and the equi-

  Auxiliary equipment (not shown) may be used to provide fuel and oxidant

  propulsion of the present invention.

  FIG. 4 shows a sectional view of a typical rocket shell 8 constructed to use the specific class of fuels according to the present invention. The casing 8 can be used as part of a rocket motor, part of a ramjet or part of a supersonic combustion ramjet. A

  mixture of hydrogen and hydrocarbon supplied by the reservoirs

  see Figures 4 and 20, respectively, of Figure 1, through

  their respective pipes and pumps is completed.

  The collector 50 is preferably disposed circumferentially around the nozzle of the rocket motor 8 and communicates with a cavity or channel 56 of fuel between a wall 54 and a wall 52. The walls 54 and 52 form the collar 62 of the nozzle and the chamber

  of combustion 64 of the envelope 8. The fuel channel 56

  which fuel is supplied by the collector 50, and wherein the fuel flows in the direction of the

  arrows 58, is contiguous, in the surrounding area, to the chamber

  broadly drawn by the zone 64 in Figure 4 and the neck of the nozzle shown in outline by the area 62 in this figure. In preferred embodiments, the channel 56 surrounds the entire rocket shell so that fuel is supplied to the combustion chamber from the entire periphery of that shell. The fuel channel can be any size that allows for an adequate supply of fuel

- 19 -

  the combustion chamber, this dimension being

  easily determined by the technician.

  According to the present invention, the combustion wall

  or the outer wall 54 and the inner wall 52, also defined as the gasket 52 of the nozzle wall, the gasket 52 of the neck of the nozzle and the gasket 52 of

  combustion chamber, depending on its location in the

  loppe 8, are made of silicon carbide fibers.

  The fibers or filaments of silicon carbide are woven and in the case of the inner wall 52 conductive heat of the combustion chamber 64 and the

  collar 62 on the fuel channel 56, and therefore

  heat for the endothermic pyrolysis of the fuel in the channel 56. It is well known in the art

  silicon carbide fibers at high temperature,

  which are wound primarily in a circumferential direction so as to form the high pressure retention for the rocket engine. Silicon carbide fibers can operate without cooling down to

a temperature of about 1200 ° C.

  In preferred embodiments of the present invention, the inner wall 52, namely the nozzle neck liner and the combustion chamber liner, is made of woven silicon carbide fibers having a porosity that selectively controls the diffusion of hydrogen between the channel 56 and the combustion chamber 64 and the neck 62, via this wall. The outer wall 54 of the channel 56 is made of high temperature silicon carbide woven fibers and is non-porous so that the hydrogen will not diffuse through

  this wall. The wall 52 is porous so that the hydro-

  gene present in the channel 56 under high pressure, for example at a pressure of about 280 MPa absolute, can pass through the wall 52, or diffuse through this wall, up to the neck of the nozzle and the combustion chamber of which

- 20 -

  the pressure is lower than that of the fuel in the channel, being for example at a pressure of about 210 MPa absolute. This diffusion of hydrogen into the wall 52 to enter the internal chambers of the casing 8 provides a fluid film cooling, as is

  represented by the arrow 90 and the layer 92 in FIG.

  As illustrated in FIG. 6, which is a large-scale portion of the firebox liner 52, the hydrogen gas 90 passes through the firebox liner or nozzle neck liner to enter the combustion chamber. combustion chamber or chamber of the neck, diffusing through the porous wall 52 of silicon carbide fibers continuously wound to form a film 92 of hydrogen on the side of the wall

  next to the bedroom and the collar. The diffusion of hydro-

  gene in the wall provides cooling by transpiration

  ration as a result of the migration of hydrogen under

  in the wall as described above, and is enabled by the porosity of the wall. It is this effect that contributes to maintaining the temperature of the filling 52

  room and neck at a value lower than the

  decomposition temperature of the silicon carbide filaments, for example at maintaining the temperature of the wall at a value less than the range between

about 980 and 1200 C.

  In the manufacture of the combustion chamber lining and the nozzle neck lining for a rocket shell containing a combustion chamber, fuel injectors, a nozzle neck and a nozzle, a fuel channel in which the fuel is subjected to endothermic pyrolysis, this channel being contiguous, surrounding them, to the combustion chamber and to the neck and having means for supplying a fuel stream to the channel, carbide fibers are woven into a multitude of layers in an armor that promotes the

- 21 -

  high pressure retention for the combustion chamber and neck; and the woven silicon carbide fibers are formed into a combustion chamber and a nozzle neck. Thus, the walls 52 and 54 are constructed with filaments or fibers incorporating silicon carbide,

  preferably in layers which are woven circumferentially

  mainly in continuous armor. The filaments or fibers of silicon carbide not only promote high pressure retention for the combustion chamber and the throat of the nozzle, but are also effective for conductive removal of heat from zone 64 of the nozzle.

  combustion chamber and zone 62 of the neck of the nozzle.

  In the winding of the filaments or silicon carbide fibers, those skilled in the art can easily use conventional techniques, for example winding the fibers around a solid core, and the weave of the weave can be chosen in the purpose of conferring the necessary porosity to the wall 52 and the absence of porosity to the wall 54. It is easy to assemble the silicon carbide fibers wound and shaped to form the channel 54 having

  optimum dimensions using conventional techniques

  well known in the art concerned.

  In order to achieve the desired porosity for selectively controlling the diffusion of hydrogen into the woven silicon carbide of wall 52, various techniques well known in the art, for example vapor deposition and / or vapor phase infiltration,

  for depositing or infiltrating organometallic compounds

  in coiled and woven fibers, these compounds allowing the selective diffusion of hydrogen into the

woven silicon carbide.

  In connection with FIGS. 4 and 5, in another aspect of the present invention, a series of fins 80

  - are arranged in the channel 56 to direct circumferentially

  the fuel passing through it in the direction of the

- 22 -

  arrow 58, whereby the fuel travels a longer path in the channel 56 and therefore increases its residence time to improve the heat transfer to the fluid and / or gas and prolong the reaction time for endothermic pyrolysis fuel. Whirling vanes 80 drive the gaseous fuel of a

  rotation, for example hydrogen and hydro-

  carbide, so as to improve the heat transfer. The

  fins 80 can be in one piece with the outer wall.

  54, or be mounted individually on this wall.

  The fins 80 may be of any suitable material capable of withstanding the temperatures and pressures prevailing in the channel 56, and are preferably laminated with high temperature silicon carbide filaments or fibers. The fins 80 are arranged to rotate the fuel as it moves between the manifold 50 and the section 64 containing the combustion chamber. The fins may be disposed in the entire fuel channel or in segments thereof. However, in the preferred embodiments of the present invention, the fins 80 are placed in the

channel 56 at the right of zone 64.

  Fuel injection orifices 68, preferably in the form of circular holes in the wall 52 in the zone of the combustion chamber 64, are arranged to continue or facilitate the swirling movement of the hot gaseous fuel represented by the arrows 78 as it passes from the channel 56 to the combustion chamber. The desired number of orifices 68 may be used in the present invention. In most embodiments, the rate of injection of the hot gaseous fuel into its inlet to the combustion chamber exceeds

  about 300 m / s at a pressure drop of about 70 MPa.

  The holes 68 inject fuel into the combustion chamber in a direction that facilitates movement

- 23 -

  circumferential fuel. In the modes of

  In a preferred embodiment, in which the lining of the combustion chamber is substantially circular in a sectional view, the holes are oriented at an angle less than 90 relative to the perpendicular to the tangent of the liner. In most embodiments, this

  angle is from about 30 to about 60; How-

  any angle of value less than 90 can be used

  which is sufficient to facilitate circumferential movement

  fuel in its entry into the combustion chamber. The propulsion system of the present invention also comprises means for introducing an oxidant, such as liquid oxygen, from a pipe 72 which is supplied with oxidant from the tanks 6 shown in FIGS. 1 and 2. The oxidant is introduced into the combustion chamber via the injection holes 66 as shown in FIGS. 4 and 5. The pipe 72 is placed in the center of the combustion chamber and has a plurality of holes or orifices 66. for the introduction of

  the oxidizer in the center of the chamber. The oxidant is preferably

  in the room in a direction that favors

  makes its circumferential movement as it enters from the conduit 72 so that the fuel and the oxidant mix in the chamber before and during combustion. In preferred embodiments, the oxidant introduction means into the chamber is substantially circular in sectional view, and the injection holes or ports 66 are oriented at an angle.

  less than 90 ° from the perpendicular to the tan-

  the means for introducing the oxidant into the chamber.

  In preferred embodiments, the angle of the oxidant injection ports is from about 30 to about 60; however, any angle

  facilitates the circumferential movement of the oxy-

- 24 -

  during its entry into the combustion chamber.

  In some preferred embodiments of the present invention, the propulsion system has an alignment of the fuel injection ports 66 and oxidant injection ports 68 to promote the overlap of fuel and oxidant the entry into the chamber, ie the fuel injection ports and the oxidant injection ports are staggered to facilitate this overlap and further promote their mixing before combustion and during of it. In this way, the fuel injected through the orifices 68 and the oxidant introduced through the orifices 66 are in layers, so that the penetration occurs between the fuel and the oxidant before and during the combustion. In this way, a high state of mixing and rapid combustion is achieved upon ignition.

of the mixture.

  An oxidant injection pipe 72 comprises a porous sleeve preferably made of fibers woven with

  silicon carbide. Hydrogen 70 is introduced by the

  74. The porous pipe 74 furnaces

  Also known is the transpiration cooling discussed above for the gasket 52 of the combustion chamber and the neck of the nozzle. Hydrogen 70 introduced through injection ports 76 mixes with the oxidant to initiate combustion and cool the oxygen injection line. In certain embodiments of the present invention, the circumferential movement of the fuel from its injection holes in the lining of the combustion chamber takes place substantially in the

  sense of the needles of a watch, and the circumferential movement

  The oxidant entering the combustion chamber through its injection holes is practically anti-clockwise. In

- 25 -

  other embodiments, the circumferential movement of the fuel in the injection holes formed in the

  filling of the combustion chamber is effec-

  anti-clockwise movement, and the circumferential movement of the oxidant in its injection holes of the introducing means takes place substantially in the

clockwise.

  According to the present invention, there is also provided a propulsion system which uses a hydrocarbon as a fuel in which means are used to obtain a high heat flow to the chamber lining.

  combustion and that of the neck of the nozzle, the heat

  of the combustion in the casing of the rocket or combustion chamber and the throat of the nozzle, which results in the temperature of the linings exceeding their thermal limits, as well as means for cooling the lining of the chamber of combustion and that of the neck of the nozzle by endothermic pyrolysis of the hydrocarbon in the fuel channel. Thus, in addition to convective and fluid film cooling by fuel as discussed above, the hydrocarbon is decomposed, cracked or dehydrogenated by

  endothermic pyrolysis in the channel 56 before combustion.

  With the use of ethylene (C2H4) as an example, the following reaction occurs in the channel in the presence of heat and in the absence of catalyst:

2C2H4 -> C4H8

Heat C4Hg8> H2 + C4g6 Heat

H2 + C4H6 -> 2H2 + 2C2H2

  Heat The previous endothermic reaction gives 0.9 kg of gaseous hydrogen and 11.7 kg of gaseous acethylene (C2H2)

  from 12.6 kg of ethylene (C2H4). The heat absorbs

- 26 -

  during the reaction is 2558 kJ per 0.45 kg of ethylene.

  Because of this extremely high combustion temperature, acethylene is not less than hydrogen

  terms of maximum rocket thrust and specific impulse

  fic. Thus, according to the present invention, the acethylene is

  the product resulting from the endothermic pyrolysis of

carbides that are preferred.

  As seen above, when the hydrocarbon is ethylene, the products of the endothermic pyrolysis include acethylene and hydrogen. When the hydrocarbon is ethane, the products of the endothermic pyrolysis include methane, acetylene and

  hydrogen and traces of various other hydrocarbons

  bures. When the hydrocarbon is propylene, the products of endothermic pyrolysis include methane, ethane, ethylene, acethylene, and hydrogen and traces of other hydrocarbons. When the hydrocarbon is propane, the products of the endothermic pyrolysis include methane, ethane, ethylene, acethylene, propylene and hydrogen and traces of various other hydrocarbons. Naturally, it remains within the scope of the present invention to use various mixtures of the foregoing hydrocarbons. In any case, the hydrocarbons used in the present invention are those whose endothermic pyrolysis provides a product with velocities.

  of higher combustion, a product with

  higher combustion rates and / or a product having a lower molecular weight than the hydrocarbon from which it is derived. As indicated above, the products are formed in the fuel channel 56 in the presence of the heat developed in the combustion chamber and the neck of the nozzle, heat which is conducted via

wall 52.

  A catalyst can be used in conjunction with

  hydrocarbon to accelerate the change of composi-

- 27 -

  this hydrocarbon, thereby providing cooling

  Further endothermic wall 52 and, in some cases, for example in that of acethylene, endothermic pyrolysis gives a higher product for the fuel rocket. For example, the means of catalysis

  for endothermic pyrolysis can be a catalytic bed

  84 is located in the fuel channel 56 as shown in FIG. 5, the catalyst 84 is illustrated in the form of beads. The catalyst can also be applied to a wall of the channel 56, for example on the wall 54 and / or the wall 52. The catalyst can also be incorporated in the hydrocarbon so that, during the propulsion of the hydrocarbon in the channel 56, the catalyst in the hydrocarbon accelerates its endothermic pyrolysis. In the present invention, any typical catalyst well known in the art of endothermic pyrolysis and cracking, including catalysts such as

platinum and palladium.

  In the previous process, the energy is ex-

  treats the combustion chamber and the throat of the nozzle via the wall 52 to be transmitted to the fuel in the channel 56, then is returned to the combustion chamber by means of the orifices 68 for fuel injection . There is no loss or gain of thermal energy; however, the wall 52 is cooled

  by the endothermic reaction, and temperatures are

  tures and combustion products more favorable, which leads to a higher speed in the nozzle and a

greater thrust.

  Accordingly, the present invention provides a

  process for improving the fuel system of a system

  propulsion system comprising a rocket casing with a combustion chamber provided with a lining and an exhaust nozzle neck having a lining with a fuel channel adjacent to the liner,

- 28 -

  surrounding environment, in which the hydrocarbon

  tible crosses the canal; the heat due to the combustion of the

  fuel in the chamber is supplied to the channel by

  nally through the firebox and neck fittings; and the hydrocarbon is heated to a

  sufficient temperature to cause its endothermic pyrolysis

  in the channel, hence the heat extracted

  of the combustion chamber through the

  It reduces the temperature in the chamber to the fillings so that their thermal limits are not exceeded.

  In addition to the speed of circulation of

  which determines the combustion pressure (neck of the

  rocket strangled at sonic velocity) and the ratio of

  In order to determine the combustion temperature, the ratio of hydrogen to hydrocarbon can be varied to control the structural temperature (in addition to rocket thrust). Excess hydrogen increases fluid film cooling, reduces the combustion temperature (a rich fuel reduces the radiation of carbon compounds (combustible hydrocarbon) and promotes

  formation of relatively low molecular weight water.

  Those skilled in the art can adjust the ratio of hydrogen to hydrocarbon so as to obtain the most desirable combination of the foregoing variables; however, in

  In most cases, it is found desirable to use

  a ratio of 50 per cent by weight of hydrogen and

percent by weight of hydrocarbon.

  The present invention provides for the systematic use of

  of hydrocarbon and hydro-fuel fuels

  gene, and as such, there is no restriction as to the means by which the hydrocarbon is decomposed or transformed

  in another form of hydrocarbon. For example, the transformation

  The addition of ethylene or ethane to ethylene is well known in the art. However, such a conversion implies

- 29 -

  typically difficulties in obtaining a maximum conversion yield (a yield of about 1% by weight of ethylene from ethylene) in a

  short time (less than 5 milliseconds) to the time

  high temperatures. In addition, the conversion process, whether pyrolytic and / or catalytic, can form undesirable carbonaceous products, such as carbon-carbon molecules (coke) or a metal carbide (from the catalyst). A particularly undesirable reaction is

  the detonation of aceto as a result of the polymerization

carbon-carbon.

  In one aspect of the present invention,

  hydrocarbons are transformed into more

  rable, that is, fuels having higher combustion rates, higher combustion temperatures and / or lower molecular weights, by thermal transformation at a temperature of from about 920 to about 1200 C using the hydrogen already available in the fuel mixture as supplied

  through the collector 50, in order to accelerate a molecular reaction

  in which a single molecule of hydrogen is

  It is based on hydrocarbon and combines with two hydrogen atoms from the hydrocarbon to form a more desirable product, such as acethylene. For example, at a temperature of about 920 C, ethylene in

  presence of hydrogen forms of acethylene. The energy of

  tivation is reduced, and the speed of conversion accelerated.

  In addition, the excess hydrogen buffers the formation of carbon carbon molecules including the polymerization of carbon because the molecules of hydrogen are placed in the free path of the carbon free atoms, where the blocking of a carbon atom to another carbon atom. Naturally, an additional means of catalysis can also be used with the cooling means of the lining of the combustion chamber and that of the

26 1 0 0 4 4

- 30 -

  neck of the nozzle by endothermic pyrolysis of the hydrocarbon

  in the presence of hydrogen in the fuel channel, whereby the hydrogen of the fuel accelerates the rate of endothermic pyrolysis. When the fuel is a hydrocarbon, whose endothermic pyrolysis results in the formation of coke and carbon polymers, the hydrogen of the fuel also prevents the formation of coke and the polymerization of carbon, which avoids

  the detonation in the fluid channel. We can also use

  in the present propulsion system, the conventional catalyst means discussed above. Although it is not desired to limit endothermic pyrolysis at a particular temperature, the hydrocarbon and hydrogen in the fuel channel are heated to a temperature of from about 920 ° C to about 1200 ° C. storage tank either in gaseous or liquid form or a mixture of both, and although the hydrogen of the storage tank is a liquid or a dough, the hydrocarbon and hydrogen present in the channel

  56 are usually in the gaseous state because of the temperature

  high erosion prevailing in this channel.

  Although not shown, those skilled in the art can easily provide a means for igniting the fuel in the presence of an oxidant in the combustion chamber. For example, the ignition means may be an electrical system

  providing an electric arc in the area of the ignition system

  fuel injection, for example, into that of the hydrogen injection ports 76, orifices 68 for injecting

  fuel and oxidant injection ports 66.

  The preceding features, taken alone or

  in combination, provide systems and procedures for

  advanced propulsion dies using hydrocarbon

  bures or hydrocarbons in combination with hydrogen

  in the presence of an oxidant. Previous systems provided

  also have sophisticated aircraft that include

- 31 -

  production systems using a dual fuel system and a method of storing fuel in a vehicle having a propulsion system that utilizes a

dual fuel system.

- 32 -

Claims (13)

  A propulsion system comprising a rocket shell (8) containing a combustion chamber (64), fuel injectors (66), a nozzle neck (62) and a   nozzle, a fuel channel (56) in which fuel   It is subjected to endothermic pyrolysis, the fuel channel being contiguous, surrounding them, to the combustion chamber and the neck of the nozzle; and means (50) for supplying a fuel stream to the channel, characterized in that it comprises an inner wall (52) woven of silicon carbide fibers which form a lining of the combustion chamber and a neck lining. the nozzle and an outer wall (54) woven of silicon carbide fibers spaced from the inner wall to form the fuel channel, whereby the silicon carbide fibers are woven so as to conduct the heat between the combustion chamber and the neck of the nozzle and the fuel channel, hence the provision of heat for the   endothermic pyrolysis of the fuel.   Propulsion system according to claim 1,   characterized in that it further comprises an integral wall   silicon carbide fiber woven fabric having a porosity for selective control of hydrogen diffusion between the channel, via the wall,   and the combustion chamber and the neck of the nozzle.   Propulsion system according to claim 1, characterized in that the silicon carbide fibers are wound substantially in a circumferential direction to provide a high pressure containment vessel.   for the combustion chamber and the nozzle neck.   4. Propulsion system according to claim 1, characterized in that the inner wall woven carbide fibers is maintained at a temperature between a value of less than about 980 C and a temperature of about 1200 C directing the heat from the combustion of the fuel in the combustion chamber, through the inner wall, to the fuel in the fuel channel, resulting in the combustion chamber lining and the lining of the fuel neck of the nozzle are cooled by conduction of heat between the combustion chamber and the neck of the nozzle and the channel. 5. Propulsion system according to claim 2, characterized in that the silicon carbide fibers are wound substantially in a circumferential direction to form a high-pressure containment vessel for   the combustion chamber and the neck of the nozzle.   6. propulsion system according to claim 2, characterized in that the woven fibers of silicon carbide of the inner wall are maintained at a temperature between a value of less than about 980 C and a temperature of about 1200 C by conduction of the heat of the combustion chamber and the neck of the nozzle   through the inner wall, towards the combus-   tible while in the channel, while simultaneously allowing the diffusion of hydrogen in the combustion chamber and the nozzle neck to ensure the cooling of the lining of the combustion chamber and the lining of the neck of the nozzle.   A method for producing a combustion chamber liner and a nozzle throat liner for a rocket envelope containing a combustion chamber, fuel injectors, a nozzle throat and a nozzle; a fuel channel in which fuel is subjected to endothermic pyrolysis, the fluid channel being contiguous, surrounding them, to the combustion chamber and the neck of the nozzle; and having means for supplying a fuel stream to the channel, characterized in that it comprises: - 34 -   a) weaving silicon carbide fibers in a weave which facilitates high pressure retention for the combustion chamber and the neck of the nozzle; and b) shaping the woven silicon carbide fibers into the shape of a combustion chamber and a nozzle neck.   8. The method of claim 7, characterized in that it further comprises weaving silicon carbide fibers to confer a porosity allowing the selective control of the diffusion of hydrogen in woven silicon carbide.   9. The method of claim 7, further comprising treating the woven silicon carbide fibers with a chemical vapor and depositing the vapor on the woven fibers to impart porosity to selectively control the hydrogen diffusion in woven silicon carbide.   The method of claim 9, characterized in that   that chemical vapor is an organometallic compound.   11. Process according to claim 10, characterized in that the organometallic compound is chemically deposited. in the vapor phase.   12. Process according to claim 10, characterized in   what the organometallic compound is deposited by infiltrating chemical vapor phase.   13. The method of claim 7, characterized in that the silicon carbide fibers are substantially   woven in a circumferential direction.