WO2006001927A2

WO2006001927A2 - Carnot cycle jet and rocket engine configurations

Info

Publication number: WO2006001927A2
Application number: PCT/US2005/016881
Authority: WO
Inventors: Marius A. Paul
Original assignee: Marius Paul A
Priority date: 2004-05-17
Filing date: 2005-05-14
Publication date: 2006-01-05
Also published as: WO2006001927A3

Abstract

Carnot cycle aircraft engines having a housing having a main air intake and a common discharge nozzle, a central rotor shaft with a central compressor unit supported by struts in the main air intake wherein the compressor unit has a fan unit with hollow fan blades and an air intake to the hollow fan blades and an annular ducted combustion chamber in the housing wherein air compressed in the hollow fan blades is ejected into the combustion chamber wherein the combustion chamber includes a cascade combustor with a series of fuel injectors for isothermal combustion with combustion gases discharged in an annular expansion nozzle with a variable geometry control before being discharged into the common discharge nozzle with air that bypasses the compressor unit and flows between the fan blades to the common discharge nozzle.

Description

CARNOT CYCLE JET AND ROCKET ENGINE CONFIGURATIONS

RELATED APPLICATIONS This application claims the benefit of the following provisional applications: MULTIPLE STAGES INTEGRATED ISOTHERMAL ROCKET SYSTEM, U.S. Provisional Application 60/572,175, filed May 17, 2004; NUCLEAR TURBO PROPULSION SYSTEMS, U.S. Provisional Application No. 60/572,144, filed May 17, 2004; and UNIVERSAL TURBOFAN, TURBOJET, AND AIR ROCKET SYSTEMS, U.S. Provisional Application No. 60/572,026, filed May 17, 2004.

BACKGROUND OF THE INVENTION This invention relates to the use of the Carnot cycle in aircraft engine designs to improve performance by isothermal compression systems and isothermal combustion systems. Typically, where rocket propulsion is utilized, the engines are rarely reused, the rocket engine being discarded after a single flight. Rocket propulsion is desirable for high altitude and space flight, where the level of oxygen in the atmosphere is inadequate to support conventional combustion. However, general rocket technology relies on multiple stages of rocket engines to propel a payload through the atmosphere to high altitudes where rocket propulsion is required. The cost of multiple stages of lost rockets is related to the number of stages and the size of the rockets. In rocket technology, the power density is typically limited to the maximum level of combusted gas mass flow through the minimum cross section of the venturi nozzle at the speed of sound. The practical limitation to maximizing mass flow is associated with high thermal and mechanical stresses in the hot combustion chamber and venturi reaction nozzle. These practical limitations are a deterrant to developing reusable rocket engines. For similar reasons, rocket propulsion is typically not incorporated into air breathing engines as a booster or for high altitude flight. Generally, rocket engines are only used for military projects with their high cost and single flight performance eliminating rocket technology for commercial application. Additionally, use of conventional jet engines for high altitude flight has major drawbacks. High performance jet engines are created with specific structural and functional limitations to adapt the engine to specialized flight at subsonic or supersonic atmospheric flight or high altitude flight. The difficulty in designing a universal aircraft engine for all flight conditions is in part due to the reliance on the Brayton cycle which results in a compromise between efficiency and power. With turbine temperatures limiting the efficiency of Brayton cycle engines, maximum efficiency of 30% is a result of existing compression technologies result in a maximum efficiency of 30%. At high polytropic-adiabatic compression ratios, the temperature of the compressed air is raised and the result is a reduced combustion and power capability. At the limit, when the temperature of compression approaches the temperature of combustion, the power is zero. Typically, metallurgical considerations limit high temperature turbines to a maximum temperature of 1200⁰C, thereby limiting pressure ratios to 25/1 and efficiency to 30%. At these levels, air fuel ratios are 60/1, which is four times higher than the stoichiometric ratio of 15/1. When using only 25% of the air for combustion, the remaining air is used to limit excessive temperatures. The final result is an engine four times larger than necessary or an engine that generates only a quarter of the potential power. In a similar manner, when nuclear thermal sources have been used for aviation propulsion systems, the results have been disappointing. Again, the main drawbacks of the Brayton cycle come into play. The inherent interrelation between temperature, pressure ratio, efficiency and power density create a conflict. The thermal efficiency of the Brayton cycle is directly related to the highest pressure ratio possible. Unfortunately, to achieve a high pressure ratio, the temperature of the compressed air wil be high, thereby limiting the ability of the nuclear thermal source to transfer heat to the propulsion medium. Again, when the temperature of compression approaches the temperature of the thermal source, the heat transfer will be zero. The Brayton cycle with conventional adiabatie-polytropic compression is therefore inappropriate for nuclear propulsion. These and other problems, resulting from engine design relying on the Brayton cycle, inspired investigation of other thermodynamic cycles and the discovery of the advantages of implementing engine designs using the Carnot cycle, particularly for aircraft engines designed for high speed, high altitude or simply high efficiency operation. SUMMARY OF THE INVENTION The aircraft engines of this invention avoid the inherent limitations of the Brayton cycle engines by adopting strategies that incorporate the Carnot cycle in either the combustion process, the compression process, or both. In the combustion process, as demonstrted by the rocket engine embodiment disclosed in this application, a fundamentally new concept is adopted, wherein the power of the rocket is the summa of a cascade of multiple stages of successively larger, concentric isothermal combustors and serial ejectors. In principle, an unlimited number of flow sections of growing geometrical size can be assembled to generate the thrust desired. The new rocket structure is formed with a precombustion chamber, followed by an expanding cascade of multiple isothermal concentric combustors and ejectors in a conical configuration with a final adiabatic reactive nozzle. In the embodiment of a basic rocket engine, the system is preferably contained in a pressurized cryogenic oxygen enclosure under the pressure of the evaporating liquid oxygen supplied to the combustion process. The low temperature encasement dramatically reduces the thermal stresses in the active components of the rocket engine. Liquid fuel is in part injected into the precombustion chamber with the oxygen gas, and the remaining fuel is dispersed on the hot surfaces of the successive ejectors for a regenerative endothermic process of evaporation. The mixing of fuel vapors with the cryogenic oxygen gas in the core stream of central combustion produces an isothermal combustion and expansion until a final adiabatic expansion and reaction in the discharge nozzle. The concepts of the isothermal rocket combustion are adapted for inclusion in hybrd turbojet and turbofan engines which include isothermal compression for greatly improved efficiency. In designing a high speed, high altitude aircraft engine, the environment is advantageous to switch from the Brayton cycle to the Carnot cycle. Substantial efficiencies can be achieved by utilizing the cold, high altitude air to chill the compression process and deliver cold, dense air to the thermal source. The concepts disclosed enable universal engine configurations to be constructed that include air breathing rocket propulsion that can be supplemented with oxygen for high altitude and space flight. Additionally, by controlled combustion, the engines are capable of stoichiometric combustion with air compression generated using specially cooled turbine blades and ram driving turbofans. Certain of the engine embodiments are designed to be converted from turbofan, turbojets and air rockets for subsonic, supersonic and air rocket propulsion, and with the addition of oxygen for space propulsion. Using the concepts and techniques described, the propulsion systems can achieve compression ratios over 100/1 and operate at the temperatures of stoichiometric combustion. In addition, using the concepts described, the engine propulsion systems can be applied to nuclear aviation propulsion. By adopting the Carnot cycle, use of a passive heat source as a heat exchanger is viable because the highly compressed air is maintained at a low temperature by isothermal compression. When the delta T or difference in temperature between the compressed motive gas and the heat source of the heat exchanger is high, thermal energy can be efficiently tranferred to the motive gas to enable reactive expansion. Furthermore, combining an active fuel combustion cycle to the thermally heated air can boost the temperature to stoichiometric levels without thermal damage by using the isothermal combustion techniques described for the rocket engine. The systems described have application in bi-dimensional propulsion in vertical take off and landing craft (VTOL), wherein the combined thrust can be integrated over and through the wing in producing a universal mobility vehicle with relatively unlimited range and flying capabilities. These and other features of the invention will become apparent from a detailed consideration of the Preferred Embodiments described in this specification. BRIEF DESCRIPTION OF THE DRAWINGS

Fig. 1 is a schematic of a rocket engine in cross section illustrating isothermal combustion. Fig. 2 is a schematic of a turbojet air rocket engine illustrating isothermal compression and combustion. Fig. 3 is a schematic of a turbofan air rocket engine in cross section illustrating isothermal compression and combustion. Fig.4 is a schematic of a two-stage turbojet air rocket engine in cross section illustrating isothermal compression and combustion. Fig.5 is a schematic of an alternate two-stage turbojet air rocket engine in cross section illustrating isothermal compression and combustion. Fig.6 is a schematic of a counter-rotating, two-stage turbojet, turbofan air rocket engine in half cross section illustrating isothermal compression and combustion. Fig.7 is a schematic of a two-stage turbojet air rocket engine in half cross section illustrating isothermal compression and isothermal combustion with triple flow. Fig. 8 is a schematic of a turbofan, turbojet engine with a nuclear heat source shown in half cross section illustrating isothermal compression and partial isothermal combustion. Fig.9 is an abbreviated schematic of the engine of Fig. 8 with an added nuclear heat source shown in half cross section illustrating isothermal compression and partial isothermal combustion. Fig. 10 is a schematic of a turbofan turbojet engine with nuclear heat sources shown in cross section illustrating isothermal compression. Fig. HA is a schematic of the engine of Fig. 10 in cross sectional profile illustrating bi-dimensional thrust. Fig. HB is a schematic of the engine of Fig. Ha in partial top cross section illustrating installation in a wing structure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to the engine embodiment of Fig. 1, a rocket engine having multiple stages of isothermal combustion is schematically illustrated. The rocket engine 10 comprises a high pressure enclosure 12 with a precombustor 16, a primary first venturi nozzle 18, an isothermal cascade combustor 20 and a final adiabatic reaction nozzle 22. Between the outer housing 14 and inner combustors 16, 20 and nozzles 18, 22, is a plenum 24 that provides a pressurized cryogenic oxygen tank 26. During operation, the pressures and thermal stresses in the combustors 16, 20 and nozzles 18, 22 are reduced by the pressure and cooling effect of the gasifying oxygen in the plenum 24. The precombustor 16 has a nearly spherical precombustion chamber 28 with at least one oxygen injector 30 and at least one fuel injector 32. The cascade combustor 20 is an expanding series of stepped, concentric ejectors 34 with fuel injectors 36 that direct cooling fuel along the cascade combustor 20. The mass of isothermal combustion gases expand to the final adiabatic reaction nozzle 22, where the gases are ejected from the engine for propulsion. The rocket engine 10 has three functional zones schematically indicated in Fig. 1: an oxygen supply and precombustion zone A; an isothermal combustion and expansion zone B, and; an adiabatic expansion zone C. The three functional zones create a new thermodynamic rocket cycle with virtually unlimited power density and a maximum efficiency. The protected thermal and mechanical structure of the rocket engine provides reliability and re¬ usability, eliminating the high cost of conventional disposable rocket systems. Referring now to Fig. 2, a turbojet air rocket engine 50 is shown that operates on the high efficiency Carnot cycle. The turbojet air rocket engine 50 has a narrow profile housing 51 that enables the engine 50 to be installed in an aircraft wing 52 as illustrated. The engine 50 has an air intake 53 with central rotor shaft 54 supported in part by struts 56 and in part by a central compressor housing 56, which in turn is supported in the center of the main air intake 53 by struts 60. The central compressor housing 56 has a core air intake 61 to a multistage axial compressor 62 with alternating sets of rotor blades 64 and counter-rotating blades 66, driven by a planetary gear box 67 for initial compression of the intake air. The compressed intake air is then delivered to the intake 68 of a hollow turbofan unit 70 for isothermal centrifugal compression. Air from the main air intake 53 that bypasses the core air intake 61 for the axial compressor 62 and turbofan unit 70 flows around and between the hollow fan blades 72 of the turbofan unit 70 to cool the compressed air within the fan blades 72. This bypass air then flows through the engine 50 to a common discharge nozzle 74. An annular ducted combustion chamber 76 is arranged around the turbofan unit 70 such that the cooled compressed air of the fan blades 72 passes through integral turbine blades 78 at the ends of the fan blades 72 to cool the turbine blades 72 before being ejected from the open tips 80 of the fan blades 72 into the annular ducted combustion chamber 76. In the annular ducted combustion chamber 76, a part, if not all, of the air flow passes back around to a circumferential series of turbine nozzles 82 in an ignited air-fuel combustion mixture. Fuel injected by fuel injector 84 in back passage 86 provides a high temperature combustion gas to drive the internally air cooled turbine blades 78 located at the ends of the hollow fan blades 72. As noted, combustion can occur at, or near, the stoichiometric level because of the internal cooling of the blades. The expanding combustion gases are then discharged through the annular expansion nozzle 88 and finally through the common discharge nozzle 74. The driven turbofan unit 70 drives the axial compressor 62 through the planetary gear box 67. When added thrust is required, the engine can be operated in a combined turbojet and rocket mode with compressed and cooled air in the annular ducted combustion chamber 76 passing to the isothermal rocket combustor 90. The isothermal rocket combustor or cascade combustor 90 utilizes the staged series of fuel injectors 92 for injecting into an expanding conical nozzle passage of ejectors 94 to cool the passage walls 96, allowing for a high temperature core combustion before the combustion gases in the annular combustion chamber 98, controlled by the variable geometry exit 100, are released into the common discharge nozzle 74. In the common discharge nozzle 74, the high temperature combustion gases from the turbofan unit 70 and the isothermal rocket combustor 90 mix with the bypass air flowing through the fan blades, heating the volume of bypass air for additional thrust. Referring now to Fig. 3, an alternate embodiment of the engine 50 of Fig. 2 is shown without the turbojet. The air and oxygen rocket engine 110 of Fig. 3 includes similar elements as previously described with a modified fan unit 112 having hollow fan blades 114 with open tips 116 that eject axially and centrifugally compressed air directly into an annular ducted combustion chamber 118. As in the prior embodiment, the bypass air that does not enter the compression process passes around and between the fan blades to cool the internally compressed air

U used for combustion. The annular ducted combustion chamber 118 includes an isothermal rocket combustor 119 with a staged series of fuel injectors 120 to feed a supply of fuel along the inner combustor wall 122 as the combustion gases are discharged from the annular nozzle 124 controlled by the variable geometry exit 126, as previously described. The combustion gases mix with the flow-through bypass air in the common discharge nozzle 74 to heat the air for added thrust. In the embodiment of Fig. 3, the multistage axial compressor 62 is driven by the fan unit 112 which reacts from the ram air driving the hollow fan blades 114. In this manner, compression is proportional to the airspeed. Initial airspeed can be achieved by rocket boosters of the type described with reference to Fig. 1. At maximum altitude, where the density of oxygen in the high altitude air is diminished, an oxygen injector 128 injects liquid oxygen into the combustion chamber 118, which gasifies before entering the isothermal rocket combustor 119. Referring now to Fig. 4, a two-stage turbojet air rocket engine 140 utilizing the Carnot cycle for cooled compression and high temperature combustion is shown. The elements of the two-stage turbo air rocket engine 140 are numbered as in Fig. 2 except where components are modified. The two-stage turbojet air rocket engine 140 has a multistage axial compressor 62 with alternating sets of counter-rotating rotor blades 64 and 66 and a turbofan unit 142 with counter-rotating turbofan rotors 144 and 146. The turbofan rotors 144 and 146 have hollow fan blades 148, 150 with integral hollow turbine blades 152, 154 at the ends of the fan blades. The hollow turbine blades 152, 154 have open tips for ejecting axially and centrifugally compressed air that has been cooled by the air that bypasses the compression process and flows between the counter-rotating blades 148, 150 and through the engine 140 to the common discharge nozzle 74. As in the Fig. 2 embodiment, the cooled and compressed airflow into the annular ducted combustion chamber 76 flows in part to the series of turbine nozzles 82 to drive the internally cooled, counter-rotating turbine blades 152 and 154. In addition, on demand, a part of the compressed and cooled air flows to the isothermal rocket combustor 90 where added fuel from staged fuel injectors are injected into a nozzle passage 94 controlled by a variable geometry exit 100. The heated combustion gases from one or both paths enter the common discharge nozzle to mix with the bypass, flow-through air and heat the air for added thrust. It is understood that various combinations and modifications can be made without departing from the subject of this invention. Referring to Fig. 5, the two-stage turbojet air rocket engine 160 includes the components as described with reference to Fig. 4 with the counter-rotating blades of the axial compressor 62 driven by a transmission 162 connected to a ram air turbine 41. Referring now to the schematic half section of Fig. 6, a universal turbojet air rocket engine 170 is shown. The turbojet air rocket engine 170 combines the isothermic fan compression with the isothermic rocket combustion for high altitude flight and includes oxygen injection and fuel injection in the main flow through path for space propulsion. The universal turbojet air rocket engine 170 has an outer housing 171 with an air intake 172 and a common discharge nozzle 174. A rotor hub assembly 176 is supported on struts 178 and 180. The struts 178 in the intake 172 are equipped with liquid oxygen injectors 182 for injecting oxygen into the flow through path to the discharge nozzle 174. The struts 180 forward of the discharge nozzle 174 contain fuel injectors 184 for injecting fuel into the flow through path for air rocket propulsion when air is present, or oxygen rocket propulsion when air is supplemented or replaced by oxygen in high altitude flight. The air intake 172 has a perimeter air scoop 186 to provide an air passage under the outer housing 171 and over the annular ducted combustion chamber 188, the staged annular rocket combustor 190 and the common discharge nozzle 174. The envelope of cold ram air in the air passage cools the hot components and the wall 192 of the engine discharge nozzle. The rotor hub assembly 176 includes a gear box assembly 194 driven by the turbofan unit 196 that has internal fuel passages 198 in the fan blades 200 to sling fuel through the hollow integrally connected bifurcated turbine blades 202. Fuel discharged from the open tips 204 of the turbine blades 202 into the annular ducted combustion chamber 188 mixes and combusts with compressed air and flows in part back around to the turbine nozzles 206, where combustion gases react with the two stages of turbine blades 202 before expanding in the turbine discharge nozzle 208 to the common discharge nozzle 174. Fuel and compressed air are also diverted to an annular staged rocket combustor 210 with a controlled geometry exit 212 providing isothermal combustion by a series of added fuel injectors 211 to maintain thermal control until the combustion gases are discharged to the common discharge nozzle 174. The rotor hub assembly 176 includes a counter-rotating axial compressor 213 with an alternating series of first blades 214 on an inner hub 216 and second blades 216 on an outer hub 218 connected to a fan rotor 220 of a fan compressor 222. The fan compressor 222 also includes hollow fan blades 224 that are supplied precompressed air from the axial compressor 213. The centrifugally compressed air in the fan blades 224 is cooled by the bypass air that bypasses the axial compressor 213 and flows between the hollow fan blades 224, between the static guide vanes 226 and between the fuel channeled fan blades 200 of the turbofan unit 198. The isothermally cooled and highly compressed air is discharged from openings in the fan blade tips 228 into the combustion chamber plenum 230 for passage through the perforated inner chamber wall 240 to supply the isothermal rocket combustor 210 and supply the air for the combustion driving the turbine blades 202. Referring now to Fig. 7, a turbojet air rocket engine 250 has a triple fan rotor assembly 252 on a common shaft 254 for combining a turbofan, turbojet and air rocket into a universal Carnot jet engine. The turbojet air rocket engine 250 has a housing 256 with an air intake 258 and a common discharge nozzle 260. The air intake has forward struts 262 with a liquid oxygen supply line 264 and a series of oxygen injectors 265 for injecting oxygen into the flow stream when ram air is insufficient to support combustion at ultra high altitude flight. The rotor assembly 252 includes a first fan rotor 266 with an air intake passage 268 to hollow fan blades 270 having an interior passage 272 for centrifugal compression of air or oxygen. The bypass air that flows between the fan blades 270 cools the compressing air (or oxygen) and the isothermally compressed air is then adiabatically compressed before combustion. A second fan rotor 274 has fan blades 276 with an interior fuel supply line 278 and a series of fuel injectors 280 for injecting fuel into the flow through air and oxygen mix when operating in rocket mode. The ends of the fan blades 276 have a series of peripheral axial compressor blades 282 that alternate and co-act with axial compressor stator blades 284 in an annular axial compressor 286 for adiabatically compressing the isothermally precompressed air delivered from annular passage 288. A third fan rotor 290 has fan blades 292 with a fuel supply line 294 leading to a bifurcated end 296 with integral turbine blades 298, 299, which are cooled by the fuel before the vaporizing fuel is ejected from open tips 300 of the turbine blades 298 into a staged isothermal rocket combustor 302. The third fan rotor 290 is driven by combustion gases from a ducted annular combustion chamber 304 with a fuel injector 306. Part of the combustion gases pass through the two stages of turbine blades 298, 299, driving the fan rotor 290 before discharging through concentric discharge channel 308, controlled by the variable geometry exit system 310. The remaining combustion gases pass through the isothermal rocket combustor 322, supplemented by the staged cooling fuel ejected from the two stages of turbine blades 298, 299 and into the outer concentric discharge channel 312 controlled by the variable geometry exit system 314. The three reactive jet streams combine in the common discharge nozzle 260. The third fan rotor 290 is an active turbofan that drives the connected second fan rotor 274 and first fan rotor 266 at lower airspeeds and at higher speeds the three fan rotors are driven by the ram air and are converted into ram air turbines. The triple reactive propulsion gases generated from the turbofan, turbojet and air rocket provide a high velocity universal Carnot jet engine. At high altitude when the density of oxygen in the air is diminished, liquid oxygen is emitted from the injectors 265 and liquid fuels (LH2, LNG, jet petrol, etc.) are emitted from the injectors 280 into the main flow stream forming a total universal combustion and propulsion system. The Carnot cycle can be adapted to novel combined thermal source engines, including a fixed heat source from a nuclear reactor and a variable fuel combustion source with high performance results. Referring to Fig. 8, the schematic illustration of a combined nuclear and combustion engine 350 utilizing a Garnot cycle for thermal efficiency is shown in half section with an engine nacelle 352 and engine core 354 supported in part by struts 356. The combination engine 350 has an air intake 358 and a common discharge nozzle 360 for reactive propulsion. At the air intake is a variable device 362 that directs air to a counter-rotating axial and centrifugal compressor assembly 363. The prewhirl device 362 has vanes 364 that cooperate with the blades 366 of an air turbine fan 368, driving alternating axial compressor blades 370 and the first stage 371 of the counter-rotating axial and centrifugal compressor assembly 363. A second air turbine fan 374 coupled to the second stage 376 of the centrifugal compressor 378 provides the external envelope 380 with the alternating counter-rotating axial compressor blades 382, co-acting with the compressor blades 370 connected to the air turbine fan 368. The second air turbine fan 374 has fan blades 384 separated from the hollow fan blades 386 of the second stage 376 of centrifugal compressor 378 by stator vanes 388. The hollow fan blades 386 of the second stage 376 of the centrifugal compressor 378 have open tips 390 that discharge the compressed and cooled air to an air passage 392 in the nacelle 352 that returns and further cools the compressed air to the engine core 354 through hollow struts 394 separated from the hollow fan blades 386 by guide vanes 387. Flow through air that bypasses the compression process provides the isothermal compression in the second stage of the centrifugal compressor 378. The hollow struts 394 further extract additional heat initially added by adiabatic axial compression before the flow through or bypass air, driven by the three air fans, is discharged to the outer reaction nozzle 398. The cooled and compressed air from the hollow struts passes to the nuclear heat exchanger 400, providing a maximum difference in temperature for maximum expansion and efficiency, using a passive heat source. To boost the temperature, a fuel combustor 402 designed for stoichiometric combustion is included in the discharge stream of heated air. The combustor 402 has a combustion chamber 403 that is provided with two working sequences. Reverse flow guides 404 direct combustion gases through a two stage turbine 406 to a first discharge nozzle 407. The turbine 406 has blades 408 with fuel channels 410 for ejecting fuel into a direct flow chamber 411 where gases are discharged to a second discharge nozzle 412 controlled by a variable control device 414. The turbine 406 is connected to the fan rotors directly and by a gearing mechanism 416 for counter-rotation of the counter-rotating fans and compressors. A triple flow propulsion is provided by the two combustion flows and the fan driven bypass flow for discharge through the common discharge nozzle 360 at the end of the engine nacelle 352. Referring to Fig. 9, an abbreviated schematic of the combined nuclear and combustion engine 350 is shown with the nuclear heat exchanger 400 in the flow of cooled high pressure air for heating the air before final temperature boosting by the fuel combustor 402. The combustion engine 350 in Fig. 9 includes an additional or expanded nuclear heat exchanger 420 for heating the bypass air before discharge. Referring to Fig. 10, a nuclear turbo propulsion engine 430 is shown in a schematic cross section. The engine 430 has an external nacelle 432, having an intake 434 and common discharge nozzle 436 with an engine core supported in part by struts 438. The engine 430 has a front fan 440, a variable geometry prewhirl device 442, followed by an air turbine 444, driving internal axial compressor blades 446 and concentric outer axial counter-rotating blades 448 driven by the central shaft 450. Compressed air is conducted through two radial struts 452, forming an intensive air cooler before being introducted into the nuclear device or heat exchanger 454 which is preferably formed from multiple heat exchanger channels 456 of Hafnium 178, activated by an auxiliary X-ray source 458, producing an intensive gamma radiation and an intensive heat that transferred to the supercooled isothermally compressed air. The heated compressed air in the heat exchanger 454 expands and drives the turbine 460, finally producing the reactive jet associated with bypass air in the flow through discharge nozzle 462. At very high speed of the flight, the compression effect of the ram air drives the front fan 440 in the manner of an air turbine which drives the central shaft 450, reducing the power requirement of the turbine 460. The excess power is diverted by opening flaps 464 with the heated air bypassing the turbine 460 and directly entering main propulsion flow in the bypass air discharge nozzle 462. Referring to Fig. HA and HB, the engine 430 described in Fig. 10 is arranged to produce a combined axial propulsion and vertical takeoff and landing VTOL configuration in cooperation with a vertical counter-rotating ducted turbofan 464 which is activated by the variable geometry gas diverter device 466. The diverter device 466 diverts a partial flow 468 from the main reactive axial gas flow 470. The vertical counter-rotating turbofan 464 is provided with peripheral turbine blades 472 and concentric inner fan blades 474. Both flows of gases and fan air are associated in a combined air/gas ratio of 25/1 which produces a significant temperature drop for the vertical flow. Preferably, the engine 430 is integrated into the structure of a wing 476. While, in the foregoing, embodiments of the present invention have been set forth in considerable detail for the purposes of making a complete disclosure of the invention, it may be apparent to those of skill in the art that numerous changes may be made in such detail without departing from the spirit and principles of the invention.

Claims

CABNOT CYCLE JET AND ROCKET ENGINE CONFIGURATIONS

CLAIMS 1. A Carnot cycle aircraft engine comprising: an outer enclosure having a partially spherical portion and a partially conical portion; a precombustion chamber in the spherical portion with a liquid oxygen supply and a fuel supply; an isothermal cascade combustor in a part of the conical portion that communicates with the precombustion chamber; and, an adiabatic expansion nozzle in the remaining part of the conical portion that communicates with the isothermal cascade combustor wherein the isothermal combustor has an expanding series of stepped concentric ejectors, each having fuel injectors wherein isothermal combustion and expansion is generated before expanding gases are adabiatically expanded in the adiabatic expansion nozzle and discharged.

2. The aircraft engine of claim 2 wherein the precombustion chamber, isothermal cascade combustor and adiabatic expansion nozzle are separated from the outer enclosure by a plenum wherein the plenum is adapted to contain cryogenic liquid oxygen.

3. A Carnot cycle aircraft engine comprising: a housing having a main air intake and a common discharge nozzle; a central rotor shaft with a central compressor unit supported by struts in the main air intake wherein the compressor unit has a fan unit with hollow fan blades and an air intake to the hollow fan blades; and, an annular ducted combustion chamber in the housing wherein air compressed in the hollow fan blades is ejected into the combustion chamber wherein the combustion chamber includes a cascade combustor with a series of fuel injectors for isothermal combustion with combustion gases discharged in an annular expansion nozzle with a variable geometry control before being discharged into the common discharge nozzle with air that bypasses the compressor unit and flows between the fan blades to the common discharge nozzle.

4. The aircraft engine of claim 3 wherein the compressor unit includes an axial precompressor;

5. The aircraft engine of claim 4 wherein the fan unit comprises a turbofan unit and the hollow fan blades have ends with turbine blades having open tips wherein the annular ducted combustion chamber is constructed to direct part of the combustion gases back around to the turbine blades to drive the turbofan unit.

6. The aircraft engine of claim 5 wherein the compressor unit has a turbofan unit with counter-rotating fan blades having counter-rotating turbine blades at the tips of the fan blades wherein a two-stage turbofan unit is provided.

7. The aircraft engine of claim 6 wherein the precompressor has counter- rotating compressor blades.

8. The aircraft engine of claim 7 wherein the central rotor shaft has a fan with fan blades at the main air intake.

9. A Carnot cycle aircraft engine comprising: an outer housing having an air intake and a common discharge nozzle; a rotor hub assembly centrally supported in the housing on struts wherein the rotor hub assembly includes an axial compressor and a centrifugal fan compressor having hollow blades wherein air precompressed in the fan compressor passes to the centrifugal fan compressor, and a turbofan unit having fan blades with fuel channels; and, an annular ducted combustion chamber having an outer plenum wherein the hollow blades of the fan compressor have open tips that discharge compressed air into the plenum, and the fan blades of the turbofan unit have turbine blades with open tips that discharge fuel into the combustion chamber wherein the annular ducted combustion chamber is constructed to direct part of the combustion gases back around to the turbine blades for discharge to a turbine discharge nozzle, and part to an isothermal combustor having a series of staged injectors for discharge to a variable control discharge nozzle, and wherein air flowing through the engine between the hollow fan blades of the fan compressor and fuel channeled fan blades of the turbofan unit pass to the common discharge unit to mix with the combustion gases.

10. The aircraft engine of claim 9 wherein the axial compressor has alternating sets of compressor blades and the centrifugal compressor has two stages wherein the turbofan unit is geared to counter-rotate the alternating sets of compressor blades and the two stages of the centrifugal compressor.

11. The aircraft engine of claim 9 wherein the struts proximate the air intake have oxygen injectors, and the struts proximate the common discharge nozzle have fuel injectors for operation of the engine in rocket mode.

12. A Carnot cycle aircraft engine comprising: an outer housing having an air intake and a common discharge nozzle; a rotor assembly having a first fan rotor with an air intake and hollow fan blades, a second fan rotor having fan blades with ends having a series of annular axial compressor blades, and a third fan rotor having fan blades with fuel supply lines and ends having turbine blades with open tips; and, an annular ducted combustion chamber wherein air compressed in the hollow fan blades of the first rotor are delivered to the annular axial compressor blades for added compression before discharge into the combustion chamber wherein the combustion chamber is constructed to deliver a part of the combustion gases back through the turbine blades and part of the combustion gases to an isothermal combustion.

13. The aircraft engine of claim 12 wherein the engine has struts centrally supporting the rotor assembly proximate the air intake wherein the struts have oxygen injectors and wherein the fan blades of the second fan rotor have fuel lines and fuel injectors wherein oxygen and fuel is injected into the bypass air flow through the engine.

14. The aircraft engine of claim 13 wherein combustion gases passing through the turbine blades are discharged through a turbine discharge channel controlled by a variable geometry exit system and wherein combustion gases discharged discharged from the isothermal combustor into an outer concentric discharge channel wherein the three discharge streams combine in the common discharge nozzle.

15. A Carnot cycle aircraft engine comprising: an outer housing having an air intake and a common discharge nozzle; an engine core with struts centrally supporting the engine core in the housing; a centrifugal compressor assembly having hollow fan blades with open tips; an outer housing passage in communication with the open tips of the fan blades; hollow struts between the housing and engine core with an internal passage in communication with the outer housing passage; and, a heat exchanger with a heat source in the engine core wherein compressed air compressed in the hollow fan blades passes to the outer housing passage and is returned through the hollow struts to the heat exchanger and heated before discharge through the common discharge nozzle with bypass air that flows over the engine core and between the fan blades before discharging in the common discharge nozzle.

16. The aircraft engine of claim 15 wherein the engine core includes an axial compressor that precompresses air delivered to the centrifugal compressor assembly.

17. The aircraft engine of claim 15 further comprising a fuel combustor that receives heated air from the heat exchanger for combustion and a turbine that is driven by gases of combustion.

18. The aircraft engine of claim 17 further comprising a heat exchanger in the flow through path between the engine core and outer housing wherein air passing between the compressed blades of the centrifugal compress and hollow struts is heated before discharge into the common discharge nozzle.

19. The aircraft engine of claim 15 wherein the heat exchanger includes a nuclear heat source.

20. The aircraft engine of claim 15 wherein the engine is installed in an aircraft wing and has two-dimensional discharge for vertical left and forward propulsion.