PROCESS OF INTENSIFICATION OF THE THERMOENERGETICAL CYCLE AND AIR JET PROPULSION ENGINES

Abstract: The process of increasing the efficiency and output of a gas-turbine having blades in a combustion chamber (7) which includes cooling the gas-turbine blades (10) with a liquid-fuel and mixing the fuel used for cooling with air in the combustion chamber (7) and burning the mixture and driving the turbine (15) by the exhaust from the combustion chamber (7). This allows the fuel and air to be in stoichiometric mixture. Preferably the fuel is a liquid-gas such as hydrogen. The turbine (15) may be cooled by injecting liquid-fuel onto its disk (11) to provide a film coating which is centrifuged around the blades whereby the liquid will cool both the interior and exterior of the blades and the absorbed heat is introduced into the combustion chamber (7). The combustion chamber (7) includes a plurality of air inlets (8). The turbine (15) drives a compressor, and the compressor and the air inlet may be cooled with the liquid-fuel, which may be cryogenically cooled. An improved jet engine which is convertible from a turbo to a ram-jet is provided.

Description: PROCESS OF " IOTENSIFICATION OF THE THERNDENmGETICAL CYCLE AND MR JET PROPULSION ENGINES Background of the Invention This invention is made up of various processes which allow the intensification at a level of ther oenergetical maximum of the cycles of the air jet propulsion engines and new types of air jet propulsion engines. These solutions allow the increase, at a level of stoichiometric maximum of the gas temperature at the turbine-entrance simultaneously with an increase of the efficiency of the compression- process under the conditions of reducing the num¬ ber of stages of compression. The solutionsallow the use of non-petro¬ leum fuels in order to extend the capacity of adap¬ tion to the world energy situation. Actual gas-turbines, working in an open cycle and without recovery of the energy of the exhausted-gases, depend on the pressures ration 7J =p2/pl and on the value of the temperature of the burnt-gases at the turbine-entracne T_ from the point of view of the efficiency and of the power. If we neglet the losses and refer to the ideal thermal cycle, the thermal cycle, the thermal efficiency is dependent of. If and the effective work of the process is chiefly dependent on the value T_. , IPO From which we may conclude that the re¬ search, now in progress in the world is directed toward the important increase of the two parameters ff and T 3 . If we take into account the efficiency of the machines implied in this process as: the compressor, the turbine, the combustion chamber as well as the mechanical losses, we obtain the effec¬ tive efficiency: and the real mechanical work: We - Wι * foe x ^ ήcom *fM ) The conclusion remains available if we particularize it in the diagram of Figure 4 where ^ 6 =0.84; ) j τ =0.88 ? OMB =0.97; M =0.99. The two parameters are connected by rela¬ tions of optimizing, which indicate optimal values for each pair 7? and T, thus: Efficiency % 19 24 28 30 32 36 t°C 600 700 800 900 1000 1200 V P 1 8-9 12-14 15-17 16-18 18-21 27 The optimal compression ratios would be at values of if =45 to 65 and in the future even higher at the limit produced by the stoichiometric combustion with gas-temperatures of 2000-2200°C. * However, at these levels, the construction and the technological difficulties could not be obtained with present technical means. The increase of the temperature of the burnt-gases up to 1300°C. at the turbine-inlet, in cruise-conditions, by the help of cooling of the blades profile by a film of air, constitutes a practical and technological limit of the actual stage of the gas-turbine evolution. The latest re¬ searches regarding the manufacturing of the turbines of monocyrstal alloy add 100°C. thereby reaching the value of 1400 C. However, ulitization of a part of the compressed-air for the blades cooling, dimi¬ nishes the gas-turbine efficiency by the energy carried away by air and by non-participation in the combustion process. For that reason, at present, it is desi¬ rable to reduce the quantity of the cooling-air and find other ways of bearing the high gas-temper- ture. The introduction of ceramic material, with the difficulties characteristic for these materials, limits this technique in the field of the increase of the gas-temperature. In parallel with the mentioned difficulties, the increase of the compression-ratios over 25-27, leads, to great construction difficulties. The increase of the number of the compressor-stages creates prohibitive levels of complexity; this fact leads to spectacular rises in gas-turbines prices and, respectively, to the decrease of capacity of these thermal machines to compete. The cooling of the blades of the gas- turbines by fluids in closed circuit implies special difficulties and an increase of the gas-turbines complexity at higher levels than those of the air cooling. The low efficiency of the jet propulsion engines, associated with the gigantic powers, achieved or able to be achieved by these engines made a gigantic petroleum consumer of the military and commercial air force. All the shown above elements unfavorably affect the position of the gas-turbine in the world energy context. Summary This invention puts an end to the above difficulties by the fact that: It assures the cooling of the blades of the gas-turbines and of all the hot surfaces by fluids and air in an open recuperative circuit, allowing the increased cooling at a thermodynamic (stoichiometric) maximum of the temperature of the burnt-gases, which pass through the turbine. The cooling fluids together with the air form rich mixtures with great convective capacity in the cooling system; the cooling fluids may be consti¬ tuted of petroleum-fuels or liquified gases of liquid hydrogen type, from liquid methane, liquid air, water, etc. The temperature of the interior surfaces, cooled by the emulsion of air and liquid is main¬ tained above or below the cracking temperature of the hydrocarbons, and in the case of the mono- component liquids (liquid hydrogen, etc) the tempera¬ ture of the walls is indifferent. Some quantity of the cooling fluid (liquid fuel, etc) is bilaterally injected on the turbine disk, where a film coating is produced, being cen¬ trifuged toward the turbine-blades gases, from where -β- the coolant, together with the cooling air enters the interior channels of the blades, producing a convective evaporating cooling, which is finalized by a rich mixture of air and fuel, centrifugally and tangentially expelled into the zone of turbu¬ lent mixing of the combustion chamber, where the formation, of the primary mixture takes place. Another quantity of the cooling fluid (fuels or liquified gases) is injected in the in¬ terior of the stator-blades, then evaporated, mixed and emulsified with the cooling air and is introduced into the chamber of turbulent combustion, where it is associated with, the mixture of air and fuel, delivered by the turbine-rotor. The rich mixture, which cooled the tur¬ bine-blades is reintroduced in the annular combus¬ tion chamber, located around the turbine, integrally recovers the absorbed heat in the air used for cooling, The combustion chamber, located at the periphery of the turbine and around it, creates a turbo-recuperative space, where the rich mixture made up of the combustible cooling fluids, emulsi¬ fied with air is expelled with the peripheral speed of the turbine-rotor such as 400m/sec, which assures a high level of peripheral turbulence, allowing the carrying out of a final stoichiometric mixture of maximal quality. The combusiton chamber of the gas-turbine is supplied with fuel from direct source, by individual injectors and by the emulsion of air and fuel, which cooled the turbine. The recirσulation of the cooling air toward the combustion chamber, having a recuperative character, and the integral utilization of the air for combustion eliminates the limits of quantitative order for the cooling-air. The air-compression takes place in the axially and radially with counter-rotary blading compressor, which is driven by counter-rotary gas- turbine, the whole air-line becoming active during the compression-process. The process of air-compression in the compressor, in the case of the cryogenically supplied gas-turbine is helped by a preliminary cooling of the air at the compressor-inlet and by a continuous and parallel cooling of the compressor, which leads to a process of izothermal compression, starting from levels of initial temperatures less than the environment temperature. The cooling in advance of the air admission and the cooling of the compressed-air in the com¬ pressor is assured by the cryogenic liquids (liquid hydrogen, liquid methane, or liquid air) , for example hydrogen at 175° C, which recuperatively absorb the heat of the admitted air and the heat of the compressed-air and transfers it to the engine-cycle as a surplus of internal energy (additional calorific power) . OMΠ The cooling in advance of the air- admission absorbs the heat of the air originated in. the environment (heat-pump) and transfers it to the engine-cycle, conferring an increased efficiency, which in special cases of power- conditions can lead to a thermal cycle with gases exhausted at the initial temperature of the environment. The recuperative utilization of the heat, transferred to the hot surfaces of the chambers of primary combustion and to the chamber of after- combustion, leads to the generation of a source of active fluid (vaporized liquid hydrogen, vaporized liquid methane, etc.), which being superheated, drives a gas-turbine which produces the energy necessary for the air-compressor, for the fuel- pumps and for other users. The recuperative generator-assemply, in the case of a gas-turbine, is associated to a double air-flow; the first, a central one, passes through the compressor? the second, a concentric one, con¬ stitutes a by-pass flow of variable section, which peripherically rounds.the compressor; the assembly of the two air-flows is unified in the zone of the primary combustion chamber, the by-pass of compressed air is peripherically concentric to the chamber of primary combustion, passes a peripheral blading of the turbine and is unified in the zone of the chamber of after-combustion with the mixture of burnt-and unburnt-gases produced by the primary chamber, which passes through the interior blades of the gas-turbine. O PI At high supersonic speeds, when the dynamic compression of the air becomes important, the by¬ pass air-flow is made available by reduction of the section of the central air-inlet and by the increase of the peripheral section of the air-inlet, the engine working in conditions of ram-jet (stato- reactor) and thus providing a convertible engine. The recuperative generator assembly, in the case of a ram-jet, is associated with a multistaged battery of ejectors, which, receiving the energy of the recuperative fluid (hydrogen, me¬ thane, vaporized and superheated liquid) , assures the absorption, the mixing and the compression of the air, which participates in the process of com¬ bustion and propulsion, the successive combustion- chamber constitute in ensemble and each of them an unlimited thermodynamic cascade, whose energy is successively utilized for the internal compression of the air and of the gases, respectively for the final propulsion, allowing the carrying out of unlimited powers. The chamber of primary excitation of the ram-jet can be supplied with air and burnt- gases from an auxiliary micro-turbine of gases or can be supplied with oxygen and fuel in regime of rocket or can constitute a rocket-chamber with solid fuel. Another advantage of the present invention is the provision of a radial-radial aligned com¬ pressor-turbine super-charging unit.forming a turbo- OMf charger with a single expansion stage and a single compression stage. Both the compressor and the turbine include counter rotating rotors which in¬ creases the output of the unit four times. •* -fitlMX OMPI -ll- Brief Description of the Drawings An example of carrying out the processes and the air jet propulsion engines is given in connection with the Figures 1 through 9, which represent, as follows: Figure 1 - The ideal thermal cycle of the actual gas-turbines; Figure 2 - The variation of the thermal efficiency, depending on the pressures ratio J[ s Pz/Pt Figure 3 - The variation of the mechanical specific work, depending on the T_ temperature and on the pressures ratio ^ Figure 4 - The variation of the effective efficiency depending on the T_ temperature and on the pressure-ratio , ? ff Figure 5 - Turbo-engine, having axial and radial counter-rotary compressor, inverted combus¬ tion chamber with turbo-radial forming of the mixture counter-rotary gas-turbines, cooled by air and liquid and free turbine with variable stator; Figure 6-Recuperative turbo-engine, con¬ vertible to ram-jet, cryogenically supplied with air-precooling; izothermal, axial-radial compressor OMPI with primary combustion-chamber; multistaged radial turbine, cooled by emulsion of air and cryogenic liquid, with variable by-pass with chamber of precombustion cryogenically cooled and with chamber of pre-combustion cryogenically cooled and with bidimensional nozzle (rectangular) with variable geometry? Figure 7 - Diagram T-S of the cycle of the convertible turbo-engine with extensible thermal cycle? Figure 8 - Ejectional ram- et with thermogas-dynamical cascade, with cryogenical- recuperative cycle? Figure 9 is an elevational view of an improved turbo-compressor, and Figure 10 is a cross-sectional view taken along the lines 10-10 of Figure 9. E ; OMH Description of the Preferred Embodiments The turbo-engine, according to the invention and to Figure 5 is made up of a gas- turbine, characterized by the fact that it is made up of an interior compressor-rotor 1, with axial blades, associated with an interior radial com¬ pressor-rotor 2, having interposed an exterior compressor-rotor 3, with axial blades, and an ex¬ terior radial compressor-rotor 4, an air-duct 5, which leads the compressed-air to the air- housing 6, which contains the annular peripheral combustion-chamber 7 , provided with exterior meri¬ dian air-ports 8, and with interior meridian air¬ ports 9. A turbine-rotor 10, is provided with blades with conventional air-cooling ducts, the blades being mounted on the turbine-disk 11, are cooled by liquids-injectors 12. The stator blades 13, the second stage gas-turbine rotor 14, which actionates the interior axial and radial compressor 1, and the turbine 10 actuates the exterior axial 3 and radial compressor 4. The turbine 15 is the final stage of using the energy. The combustion chamber -7 is provided with the primary fuel in¬ jectors 16, the rotary-assembly is supported by the supports with bearings 17, 18,.19, 20, 21, 22? the stator blades of the final turbines 23 and 24 are adjustable for the optimizing of the speed- regime. The walls 25 and 26 of the combustion- chamber are completely manufactured of perforated holes 100, which allows the air-inletby perspiration on the whole metal surface of the combustion- chamber 7, and include supplimentary air inlets 27, disposed for assuring a high quality combustion. Referring now to Figure 6, the turbo- engine, recuperative, convertible to a ram-jet is made up of a flow 27 of central air, which assures the air-compression r with variable inlet 27.1 and variable outlet 27.2, a concentrical outer flow 28, with variable inlet 28.1 and variable outlet 28.2, a zone 29 of admission and precooling of air, a cryogenic cooled compressor 30, provided with a cooling jacket 31, valves 32 for adjusting the air-inlet, and valves 33 for the air-outlet, the primary combustion chamber 34 having an adjustable air-admission, controlled by the sliding divider 47 and provided with the walls 35 and 36, cooled by fluid, from which the cryogenic fuel is atomized in the combustion chamber 34, the stator 37, bladed inside, the stator 38, bladed outside, having the blades provided with conventional channels of inner cooling, the axial gas-turbine 39, provided with in¬ terior blades 40, and concentric, exterior blades 41, having the blades provided with channels of inner cooling, a by-pass compressed flow 42, final combustion chamber 43, provided with cryogenic cooled walls 44, and bidimensional, rectangular, adjustable nozzle 45 for reaction? at low speeds the air- compression is assured by the compressor 30, with a completely open air-inlet 27.1 and a completely open air-outlet 27.2, and at high speeds the dynamic compression of the air eliminates the necessity of the central compressor, the assembly becomes ram- -W∑E l OMPI jet by the complete or the partial openings of the by-pass flow 28.1 and of the outlet 28.2, the turbine remains to deliver limited energy for the fuel-pump 46? the process of admission of the air in the compressor is helped by a process of precooling through the cryogenic fluid (liquid hydrogen, liquid methane) , which assures the decrease of the initial temperature below the environ¬ mental surroundings; the air-compression process is also helped by the continuous cooling of the compressor through cryogenic liquid which assures a process of izothermal-compression? the heat, absorbed by the cryogenic fluid during the cooling process of the admitted air, during the air-compres¬ sion, during the cooling of the combusiton chambers and of the turbine, is reintroduced in the engine cycle as a supplementary calorific addition to the own calorific power of the fuel?. By the preliminary cooling and the izo- thermal compression, a part of the engine-cycle develops below the temperature of the environment, constituting a heat-pump, which adds a surplus of efficiency to the thermal cycle, this efficiency being maximal at the power-conditions, at which the temperature of the exhausted gases can reach values equal or comparable to those of the environ¬ ment. This is important to military aircraft which are chased by heat seeking missiles. OMPI The cryogenic liquid, in atomized state or in emulsional one, is led toward the hot sur¬ faces of the interior of the stator-blades, on the side-surfaces of the rotor-blades and is eva¬ cuated by perspiration and the ports toward the air¬ flow is mixed with this and is burnt into the com¬ bustion chamber 43. The running of the gas-turbine, according to the invention and to Figure 5, is assured by counter-rotary driving of the interior compressor 1, with axial blades,-and of the interior compressor with radial blades 2 and by driving of the exterior axial compressor 3 and of the exterior radial compressor 4, which pumps the air up through the channel 5, in the air-housing 6, which contains the annular peripheral chamber 7, provided with exterior-meridian air-ports 8, and with interior- meridian air-ports 9? the combustion being assured by injecting primary fuel through the injectors 16, allows the start of the gas-turbine and its working at reduced power. At high power, which make necessary the blades-cooling, when the burnt gases surpass 1000-1200 C, the injectors 12 inject cooling liquid-fuel or other liquids which settle down on the turbine-disk in film form, radially penetrate in the interior channels of the blades of the rotor 10 and together with the cooling- air changes to an emulsion with high capacity of absorption of the heat from the blades walls. The steam from the cooling-liquid and the air that that passes through the interior channels of the blades of the turbines 10, are expelled radially and tangentially into the annular, peripheral com¬ bustion-chamber 7 , where by final mixing, com¬ pletely burn at a level of stoichiometric limit. The whole flow of the burnt gases is manufactured of perforated sheet with micro-orifices on the whole surface, which produces an air perspiration, which eliminates the material contact between the metal-walls and the burnt-gases. The running of the recuperative turbo- engine, convertible to a ram-jet, according to the invention and to Figure 6, is assured by starting of the primary combustion-chamber 34, which produces hot gases, and delivers thermal energy to the walls, passed through by the hot gases. The liquid hydro¬ gen (or other cryogenic fuel) is pumped by the pump 46 into the precooled air-zone 29, passes the com¬ pressor 30 through and cools it, assuring a process of .izothermal air-compression in the compressor, enters the cooling jacket 44 of the final combustion chamber 43, from where enters the cooling jackets 35 and 36 of the primary combustion-chamber 34, a part of it is atomized into the inner cooling channels of the stator-blades 37 and 38, another part is atomized on the disk of the gas-turbine 39, penetrates into the inner cooling channels of the blades 40 and 41, from where, by perspiration, is introduced in the central and peripheral air-circuit of by-pass. Another part of the liquid hydrogen, atomized, vaporized and overheated, enters the pri¬ mary combustion-chamber and partially burns, OMPI proportionally with the quantity of air that entered the primary combustion chamber 34, which entered the chamber being adjusted by the sliding divider 46? the surplus of unburnt hydrogen (un¬ satisfied with oxygen) passes through the blades of the gas-turbine, as a superheated fluid, reaches the final combustion-chamber 43, where is provided with oxygen and completely burns, producing the useful effect of reaction, controlled by the reaction-nozzle 45 with variable geometry. At low speeds, the air-compression is assured by the compressor 30 with a complete open air-inlet 27.1 and a complete open outlet 27.2? at high speeds, the dynamical air-compression eliminates the necessity of the central compressor, the assembly changes in ram-jet by the complete or the partial opening of the by-pass circuit 28.1 and of the outlet circuit 28.2, the turbine remaining to deliver limited energy at the actionating of the fuel-pump 46. The air-admission process in the com¬ pressor is helped by a process of precooling through cryogenic fluid (liquid hydrogen, liquidmethane, etc) , which assures the decrease of the initial air- temperature (Figure 7, position 1) below the tempera¬ ture of the environment (Figure 7, position-1) . The air-compression process is helped in paralled by the continuous cooling of the compressor with cryogenic liquid, which assures an izothermal OMPI - - compression (Figure 7, position-2) . The heat absorbed by the cryogenic fluid during the process of cooling the admitted-air, during the process of air-compression, and of the cooling of the combustion-chambers and of the tur¬ bine is reintroduced in the engine cycle as sup¬ plementary thermal addition to the own thermal power (chemical) of the fuel. By precooling and by izothermal compres¬ sion, a part of the engine cycle develops at tem¬ peratures lower than the environmental temperature, constituting a heat-pump which adds a surplus of efficiency to the thermal cycle, this efficiency becoming maximal at power-conditions, at which the temperature of the exhausted-gases (isentropic expansion from optimal T. to 1) can reach equal or comparable values with those of the environment. The running.as convertible ram-jet, according to the invention and to Figure 8, takes place by starting the excitation-chamber 48, which is closed axially by the sliding cone 49. The central burnt-gases, expelled by the excitation chamber, assure the starting of the assembly by the * successive air-absorption produced by the ejection process, assured by the multistaged ejectors 52. Simultaneously, the hydrogen, vaporized and pumped by the pump 56, passes through the interior walls of the recuperative generator 54, reaches high pa¬ rameters of temperature and pressure and is expelled O PI -2.0- by the ejection nozzle 52, carrying away the air, mixing and compressing the gas-mixture, which- burns in the chamber 53, and is expelled as reactive jet by the nozzle 57. At high speeds, the cone 49 withdraws, setting the air-access free toward the excitation- chamber 48, which thus passes from rocket-mode to the ram-jet mode. The hydrogen vapors, produced by the recuperative generator, drive the turbine 55, which in its turn drives the pump 56 and this pumps the liquid hydrogen in the recuperative journey. Another excitation variant can be achieved by filling the chamber 48 with solid fuel, which assures the period of start by combustion, following that after liberation of the combustion-chamber to be passed on to full-run load. Another variant of excitation is carried out by supplying the excitation-chamber with air and burnt-gases originated in an auxiliary icro- turbo-engine. O PI The turbocompressor with counter- rotative radial rotors is made up, according to the invention and to the Figures 9 and 10, of the cen¬ tral radial compressor first rotor 59, with an in¬ let diameter Do and the outlet diameter D, common with the central axle 60 which is assembled the first rotor of the central, radial turbine 61. The compressed air from the central rotor 59 is led into the counter-rotative periphery second rotor 62, which in its role of difuser is provided with blades from diameter D 2 to j - i , is continued with a connecting zone from D„_. to D__„ and finalized with a radial zone from D _ to D_ _.. The com¬ pressed air which came out of the counter-rotative peripheral rotor 62, enters the air-difuser of the compressor 63. The burnt-gases enter the turbine by the periphery blades of the stator 64, which in an optimum way lead their flow toward the blades of the periphery second rotor 65, manufactured similar to the compressor 62 but with inverted functions. The burnt-gases, which leave the blades of the rotor 65, are led to the central turbine 61, from where they are exhausted outside. Rotors 62 and 65 are connected together. The counter-rotative rotary motion of the two concentrical turbo-compressors, assure two compression-stages, within a single turbocompressor with a single stage. Because the relative speed between the two rotors is double the termodynamical target on each counter-rotative assembly increases four times. That is, rotors 62 and 65 rotate oppo¬ site to the rotation of rotation 59 and 61 respecti¬ vely. The advantages of the intensification of the thermoenergetical cycle of the air-jet propulsion engines are as follows: It assures for all the types of gas- turbines and ram-jets the possibility of func¬ tioning at energetical levels of thermodynamical maximum, by eliminating the barrier of maximal temperature of the gases in the energetical journey of the cycle. It assures the possibility of functioning at levels of maximal and most favorable compression- pressures with a number of pressure-stages reduced with 50%, in izothermal conditions (for hydrogen cycle) with maximal efficiency. It assures the most efficient level of operation with the lowest fuel-consumption, con¬ sequently the longest operation-range with the most reduced operation-costs. It assures the possibility of function¬ ing of the aircraft-engines and of the cruise- missiles with liquid hydrogen, therefore eilimi- nates the dependency on the petroleum sources and on the supplying bases. OMPI It assures the unlimited strategical independence of the ships of the naval air- crafts, which are supplied with liquid hydrogen from the nuclear plants of propulsion of the ship- base carriers, liberated of the necessity of being supplied with petroleum for aircraft and petroleum for naval propulsion. It assures the commercial aircraft with liquid fuel-liquid hydrogen from cheap and unlimi¬ ted nuclear source, belonging to the respective airport. The convertible turbo-ram-jet engines are able to cover with maximal efficiency the largest range of speeds, from subsonic level to the highest supersonic levels. The izothermal air-compression leads to the lowest power-consumption in the compressor and to the lowest power-requirement produced by the turbine, adequately reducing the turbine dimensions and supplying supplementary useful energy. The recovery of the heat absorbed from the precooling of the air, from the izothermal compres¬ sion of the air and from the heat absorbed from all the hot parts is carried out by its transportation by the hydrogen and its reintroduction in the engine cycle becomes an important source of increasing the general efficiency of the engine. t reduces the section of the air-inlet, adequately reducing the resistance to motion. By the great expansion of the burnt- gases from the high levels of pressure, the tem¬ perature of the exhausted-gases massively decreases, adequately reducing the infrarred signature. It- reduces the heat signature of the engine. The present invention, therefore, is well adapted to carry out the objects and attain the ends and advantages mentioned as well as others inherent therein. While presently preferred emodbi- ments of the invention have been given for the purpose of disclosure, numerous changes in details of construction and arrangement of parts will be readily apparent to those skilled in the art and which are encompassed within the spirit of the invention and the scope of the appended claims.

Claims: What is Claimed Is: 1. The process of increasing the effi¬ ciency and output of a gas-turbine having blades and a combustion chamber comprising, cooling the gas-turbine blades with a liquid-fuel and air, mixing the fuel used for cooling with air in the combustion chamber and burning the mixture, and driving the turbine by the exhaust from the combustion chamber. 2. The process of claim 1 wherein the air and fuel in the combustion chamber reaches a stoichiometric maximum. 3. The process of claim 1 wherein the fuel is a liquified gas. 4. The process of claim 3 wherein the fuel consists of a group including hydrogen, methane and methanol. 5. The process of claim wherein the turbine includes a disk with extending blades and including, injecting the liquid-fuel on the turbine- disk to provide a film coating which is centrifu¬ ged toward the blades, cooling the interior and the exterior of the blades with the cooling fuel. ^ δ tangentially expelling the cooling liquid from the blades into the combustion chamber whereby the heat absorbed from the turbine is introduced into the combustion chamber. 6. The process of claim 1 including, introducing air into the combustion cha¬ mber at a plurality of locations around the outer periphery of the chamber. 7. The process of claim 1 wherein the turbine drives an air-compressor supplying air to the combustion chamber and including, cooling the compressor with the liquid- fuel. 8. The process of claim 1 including, cooling the air with the liquid-fuel prior to cooling the turbine-blades. 9. The process of claim 1 wherein the liquid-fuel is cryogenically cooled. 10. The process of claim 1 including, providing first and second air flows to the trubine, said first air flow being through a compressor, and said second air flow by-passing said compressor, and selectively adjusting the amount of air flow between the first and second air flows. 11. The process of claim 10 including, passing the exhaust, additional air and fuel to an afterburner, cooling the exterior of the afterburner with liquid fuel. 12. The process of claim 7 including, precooling the inlet air to the compres¬ sor by cryogenic liquid-fuel for decreasing the initial temperature of the inlet air below the temperature of environment. 13. A gas turbine, characterized by the fact that it is made up of an interior compressor rotor (1) , with axial blades, associated with an interior radial compressor rotor (2) , having inter¬ posed an exterior compressor rotor (3) with axial blades, and an exterior radial compressor rotor (4) , an air duct (5) , which leads the compressed air in the air housing (6) , which contains the annular combustion chamber (7) , provided with exterior air ports (8), and with interior air ports (9), a turbine-rotor (10) , provided with blades with cooling ducts, the blades being mounted on the turbine-disk (11), and cooled by liquid injectors .(12) , a second stage gas-turbine rotor (14) , which actuates the exterior axial and radial compressor, a turbine (15) is the final stage of using, the energy? the com¬ bustion chamber (7) is provided with the primary injectors (16), the stator blades of the final tur¬ bines (23) and (24) are adjustable for the optimi¬ zing of the speed, the walls (25) and (26) of the combustion chamber include a plurality of perforated O PI -23- holes which allow additional air into the combus¬ tion chamber, and including air leading inlets (27) , disposed for assuring a high quality com¬ bustion. 14. A turbo engine, convertible into a ram-jet characterized by the fact that it is made up of a flow (27) of central air for compression, with a variable inlet (27.1) and variable outlet (27.2), a concentric flox passage (28), with vari¬ able inlet (28.1) and a variable outlet (28.2), a zone (29) of admission and precooling of air, a cryogenic cooled compressor (30) , provided with a cooling jacket (31) , valves (32) for adjusting the air inlet, and valves (33) for the air outlet, a primary combustion chamber (34) , having an adjust¬ able air admission controlled by the sliding divider (47) , and provided with the walls (35) and (36) , cooled by fluid, from which cryogenic fuel is atomized in the chamber (34), a stator (37), inside a stator (38), outside, having the blades provided with channels of inner cooling, an axial gas-turbine (39) , provided with interior blades (40) and concentric exterior blades (41) , having the blades provided with channels of inner cooling, a by-pass air compressed flow (42) , a final combustion chamber (43) , provided with cryogenic cooled walls (44) , an adjustable nozzle (45) , whereby for low speeds the air compression is provided by the compressor (30), with a completely open air inlet (27.1) and a completely open air outlet (27.2)? at high speeds, the dynamic compression of the air eliminates the necessity of the central compressor and the assembly becomes a ram-jet by the complete or by the par¬ tial -openings of the by-pass flow. 15. The apparatus of claim 14 including, means for precooling the inlet air to the compressor by cryogenic liquid-fuel for decreas¬ ing the initial temperature of the inlet air below the temperature of the environment, and means for cooling the compressor by cryogenic fuel, and means for introducing the cooling fluid into the combustion chamber as a fuel. 16. The apparatus of claim 15 including, means for cooling the turbine blades by the cryogenic liquid-fuel. 17. A ram-jet, characterized by having an excitation chamber (48) which provides a primary central jet for a plurality of staged ejectors (51) , having double walls constituting the outlet nozzles (52) of the ejectors, with combustion chamber (53), provided with a jacket (54) , which constitutes a generator, an auxiliary turbine (55) , which drives a pump (56) and the variable outlet nozzle (57) , an air-admission port (58) , whereby the liquid hydro¬ gen is pumped, during flight, into the jacket of the recuperative generator (54), where by heating and vaporization, is brought to high temperatures and pressures, and is led, a part into the excita¬ tion chamber (48) providing primary combustion and the generation of a jet of burnt gases for exciting -3 Q- the ejectors, another part to the staged ejectors (51) , then is expelled with high energy by the ports (52) of the ejectors, producing the multi- staged absorption of the air admitted into the engine by the inlet (58) , the final combustion of the mixture of the air, fuel and gases of excita¬ tion taking place in the combustion chamber (53) . 18. A turbocompressor having a radial aligned compressor and turbine in which each com¬ pressor and turbine includes a first rotor, said first rotors being connected together by a central axle, sai compressor and turbine also each includ¬ ing a second rotor, said second rotors rotating counter to said first rotors, and said second rotors being connected together. Cϊ.'.n

Inventors: PAUL MARIUS A, PAUL, MARIUS, A.
Applicants: PAUL MARIUS A, PAUL, MARIUS, A.
Application year: 1984
Main classification: F02C7
IPC codes: F01D5, F02C7, F02K7, F23R3