This invention utilizes a Carnot cycle to improve the performance of aircraft engines and is designed to utilize alternate thermal sources including fuel cells, nuclear power reactors and mixed fuel combustors for combined jet and rocket propulsion permitting extended and indefinite flight of manned and unmanned aircraft.
Actual state-of-the-art conventional aerospace propulsion systems are under many limitations. Reduced gas turbine power density, reduced thrust to weight ratios, reduced thermal efficiency and reduced cycle pressure ratios are a direct result of limited maximum turbine inlet temperature, limited by the metallurgical characteristics of the turbine blades.
Liquid fuel rocket propulsion systems based on liquid oxygen (LOX), are exclusively for space propulsion and are not applicable to the field of military and commercial air aviation because of the enormous cost of operation.
The revolutionary universal Carnot propulsion systems for turbo rocketry eliminate many technological barriers which have typically resulted in limited thermal efficiencies averaging 30% with pressure ratios of 25/1, air fuel ratios of 60/1 and thrust to weight ratios of 16/1, which generally use only 25% of the air compressed in the gas turbines.
Our revolutionary Carnot rocket propulsion system opens the capability to combine atmospheric oxygen in high altitude aviation with the addition of supplemental LOX for extended altitude range into space. The most exceptional characteristic of the described propulsion systems is the common option to use fuel combustors, fuel cells and/or nuclear energy reactors in the same propulsion system, separately or in combination.
For all alternatives, the main characteristic is the maximum absolute thermal efficiency of the Carnot thermal cycle, producing the maximum absolute range of flight for the available thermal source. For all alternatives, the final compression process is isothermal, with minimum work for compression and a minimum temperature at the end of the compression. For all alternatives, the combustion process cane achieve a maximum absolute isothermal and stoichiometric level, resulting in maximum power density and the maximum absolute thrust to weight ratio possible.
State-of-the-art gas turbines are not conserving the maximum turbine inlet temperature and pressure ratio of the cycle from full load to part loads, and are thereby losing the temperature and pressure ratios at part loads. The corresponding loss of thermal efficiency results in raising the specific fuel consumption to unacceptable high levels and severely limits the range of flight.
The revolutionary gas turbine Carnot cycle engines of this invention can operate at a constant isothermal temperature resulting in better efficiency and better specific fuel consumption at all loads of operation for extended flight.
The universal Carnot cycle engine of this invention is generally designated by the reference numeral 10. The preferred engine design of this description utilizes a novel isothermal compressor system as described in my referenced application to enable a maximized thermal difference between the compressed air available for thermal heating and the practical temperature limit for combustion as defined by the structural integrity of the turbine blades or blade segments subject to the high temperature of the motive gases after heating. Since the engine 10 of this invention is designed for high altitude flight, the entry temperature of air for compression is low, in the minus range of −50° C. to −100° C., and can be used for isothermal compression of the air not passed through the core of the engine as bypass air. With a maximized delta T or temperature difference available for thermal heating and expansion of the drive gases, multiple thermal sources can be considered in alternatives or in combinations, making for high efficiency, long distance flight. As noted, by injecting liquid oxygen into the air flow passages, aerospace flight can be achieved in any of the engine embodiments described.
Referring to
The inlet temperature for state-of-the-art gas turbines is typically maximized at 1200° C. The optimum pressure ratio for the maximum inlet temperature of 1200° C. is usually 25/1. This pressure ratio corresponds to a maximum thermal efficiency of 30%. Internal cooling of turbine blades with compressed air which reaches 800° C. permits only a constant pressure combustion developing power in the limited temperature interval of 800°-1200° C. Attempting to raise the pressure ratio for improved efficiency by an increase in the air compression results in a higher temperature of the compressed air which reduces the cooling ability of the compressed air and the temperature differential available for the combustion process with an accompanied loss in power.
At a certain point, a total loss of power will result when the compressed air temperature at the thermal source inlet is raised to the maximum thermal source temperature for the turbine blades. When this occurs, no heat can be added and/or no fuel can be burned and hence, no power can be produced.
In conclusion, with the Brayton Cycle, the thermal efficiency cannot be improved if a higher turbine inlet temperature cannot be achieved.
Referring to
As shown in
The turbofan jet engine 10.1 of
The expanding gases from the turbine blades 129 of the gas turbine fan 130 discharge from an annular nozzle 133 to mix with and heat the flow through bypass air in a final combined jet nozzle 134 at the tail end 135 of the housing 120. The gas turbine blades 129 drive the fan blades 107 to accelerate the bypass air for added thrust, and drive the connected turbofan rotor 102. The thermal source 126 may alternately comprise fuel cells or a fuel combustor.
Referring to
In addition to the gas turbine fan 130, the engine 10.2 has a second stage gas turbine fan 132 that rotates counter to the gas turbine fan 130 connected to the turbofan rotor 102. The spaced turbine blades 129 of both turbine fans 130 and 132 are separated by stator blades 141 for channeling the gas flow to the annular nozzle 133 and final jet nozzle 134. The work from the second stage gas turbine fan 132 drives the integral fan blades 111 in a rotational direction opposite to the fan blades 107 of the first gas turbine fan 130. In this manner, the turbofan jet engine is a triple turbofan operable with any thermal source.
Referring to
As noted, the heat source 75 may be one of several alternatives, such as a fuel cell, nuclear reactor, or fuel combustor. In the case of a fuel combustor with liquid fuel, added cooling can be obtained by routing some liquid fuel to the air passages 78 of the gas turbine blades 76 as taught in my earlier patent applications. Final expansion in the annular reverse flow thermal chamber ejects hot gases through the annular nozzle 134 for mixing with and heating the cold bypass air passing through the engine for final discharge from the jet nozzle 136 at the end 138 of the engine.
Referring now to
In the embodiment of
As noted, the two thermal sources 148 and 149 may comprise the same or a combination of alternate thermal sources, such as fuel cells, nuclear reactors or fuel combustors. In one preferred embodiment, where the thermal sources are nuclear reactors, the reactors may be of a simplified form as shown in
In
By use of a nuclear thermal source, high altitude flight may be maintained vertically, indefinitely.
Referring now to
The turbofan jet engine 10.5 has an outer housing 120 with struts 115 and a center shaft 117 as previously described for the previous embodiments. Again, an aerial compressor 136 with a first series of staged blades 137 rotating counter to the staged blades 138 of the free wheeling fan 139 that rotates on bearings 140 supported by struts 124. The fan blades 122 rotate counter to the hollow fan blades 104 of the turbofan compressor rotor 102.
In the engine 10.5 of
The tips 118 of the hollow blades 104 are joined to the tips 198 of the fan blades 185 of the gas turbine rotor 184 to pass compressed air from inside the hollow blades 185 to the hollow gas turbine blades 199 as previously described with reference to
The majority of the compressed air is delivered into a plenum 201 around the thermal chamber 190. The inner wall 202 is perforated in a predefined manner to allow compressed air to bleed into the chamber 190, cooling the chamber walls with the majority of compressed air entering the common zone 187 for supplying the thermal sources 188 and 189.
The reverse flow path from common zone 187 cycles through the heat source 188, which may be a combustor or a fuel cell with an electrical supply 203 from generator 204 connected to bearing 196. The hot gases flow through the blade assembly 186 of the gas turbine rotor 184. The opposite flow path passes compressed air though heat source 189 for expansion in flow path 191 to variable geometry control nozzle 192 for a discharge flow 195 in combined nozzle 205.
Referring to
In high altitude flight with limited atmospheric oxygen, the air is supplemented by a liquid oxygen injection through injectors 217. Added fuel can be injected into the bypass flow by fuel injectors 218 for rocket propulsion in addition to the rocket propulsion from nozzle 214 and jet propulsion from nozzle 212.
Referring to
The turbojet assembly 402 produces the jet propulsion effect through the reverse jet gas reaction passage 405 and the ejector 406 where a mixture of dilution air for reduction of the infrared signature and amplification of the mass of gas flow is received from the air entry plenum 407. The air rocket assembly 403 is provided with a combustion chamber 408 or nuclear reactor 408.1, and a variable geometry nozzle device 409 controlling the gas jet 410.
All three jet gas reaction effects are finally combined and mixed in the gas reactive jet 411 at ejection nozzle 412 with a maximum propulsion efficiency and a minimum exhaust temperature for minimum infrared signature.
In
The counter-rotating power gas turbine turbofan 430 comprises an assembly of fan blades 431 with a ducted peripheral crown 432 that is provided with circular air collectors 433, air admission ports 434, and hollow gas turbine blades 435. The counter-rotating power gas turbine turbofan rotor 440 is an assembly of fan blades 441 with a ducted peripheral crown 442 that is provided with the circular air collectors 443 connected with a common air cooled collector 444, and hollow gas turbine blades 445 that are supplementarily cooled by internally injected fuel 446, supplied through the radial fuel slinger duct 447.
The counter rotating power gas turbine turbofan rotor 450 includes an assembly of fan blades 451 with a ducted peripheral double crown 452 that is provided with the circular air common collector 453 connected with the common air cooled collector 444 and hollow two stages gas turbine blades 455 and 456 that are supplementarily cooled by internally injected fuel 457 and 458 supplied through the radial fuel slinger duct 459. The counter-rotating internal assembly 450 also drives the axial counter-rotating compressor 460, provided with internal axial compressor blades 461, 462, 563, 464 and the centrifugal isothermal compressor 465.
The counter-rotating external assembly 470 is also provided with counter rotating axial compressor blades 471, 472, 473, 474 and 475 and is driven by the air turbine 476 that is controlled by the variable geometry devices 477 located in the strut 478.
The isothermal centrifugal compressor 46 supplies cooled compressed air in a central plenum 480 which is conducting cooled air in the left combustion chamber 481 and/or in the left nuclear reactor 482 and into the common air cooled collector 444.
The triple thermal cycle is provided with the capacity to have isothermal compression and isothermal combustion expansion as described earlier with reference to
A variable geometry intake device 490 in association with variable geometry gas turbine stator blades 491 and the variable geometry air turbine device 477 creates a constant air pressure ratio for constant thermal efficiency from full loads to part loads and minimum specific fuel consumption in all the operational capabilities of the engine in the embodiment.
This application relies on the priority of the following provisional applications: U.S. Provisional Application Ser. No. 60/621,183, filed Oct. 21, 2004, entitled UNIVERSAL PROPULSION CARNOT CYCLES TURBO ROCKETRY; and U.S. Provisional Application Ser. No. 60/646,009, filed Jan. 21, 2005, entitled UNIVERSAL PROPULSION CARNOT CYCLES IN TURBO ROCKETRY.
Number | Date | Country | |
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60621183 | Oct 2004 | US | |
60646009 | Jan 2005 | US |